[Title 14 CFR ]
[Code of Federal Regulations (annual edition) - January 1, 2015 Edition]
[From the U.S. Government Publishing Office]



[[Page i]]

          

          Title 14

Aeronautics and Space


________________________

Parts 1 to 59

                         Revised as of January 1, 2015

          Containing a codification of documents of general 
          applicability and future effect

          As of January 1, 2015
                    Published by the Office of the Federal Register 
                    National Archives and Records Administration as a 
                    Special Edition of the Federal Register

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                            Table of Contents



                                                                    Page
  Explanation.................................................       v

  Title 14:
          Chapter I--Federal Aviation Administration, 
          Department of Transportation                               3
  Finding Aids:
      Table of CFR Titles and Chapters........................    1061
      Alphabetical List of Agencies Appearing in the CFR......    1081
      List of CFR Sections Affected...........................    1091

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                     ----------------------------

                     Cite this Code: CFR
                     To cite the regulations in 
                       this volume use title, 
                       part and section number. 
                       Thus, 14 CFR 1.1 refers to 
                       title 14, part 1, section 
                       1.

                     ----------------------------

[[Page v]]



                               EXPLANATION

    The Code of Federal Regulations is a codification of the general and 
permanent rules published in the Federal Register by the Executive 
departments and agencies of the Federal Government. The Code is divided 
into 50 titles which represent broad areas subject to Federal 
regulation. Each title is divided into chapters which usually bear the 
name of the issuing agency. Each chapter is further subdivided into 
parts covering specific regulatory areas.
    Each volume of the Code is revised at least once each calendar year 
and issued on a quarterly basis approximately as follows:

Title 1 through Title 16.................................as of January 1
Title 17 through Title 27..................................as of April 1
Title 28 through Title 41...................................as of July 1
Title 42 through Title 50................................as of October 1

    The appropriate revision date is printed on the cover of each 
volume.

LEGAL STATUS

    The contents of the Federal Register are required to be judicially 
noticed (44 U.S.C. 1507). The Code of Federal Regulations is prima facie 
evidence of the text of the original documents (44 U.S.C. 1510).

HOW TO USE THE CODE OF FEDERAL REGULATIONS

    The Code of Federal Regulations is kept up to date by the individual 
issues of the Federal Register. These two publications must be used 
together to determine the latest version of any given rule.
    To determine whether a Code volume has been amended since its 
revision date (in this case, January 1, 2015), consult the ``List of CFR 
Sections Affected (LSA),'' which is issued monthly, and the ``Cumulative 
List of Parts Affected,'' which appears in the Reader Aids section of 
the daily Federal Register. These two lists will identify the Federal 
Register page number of the latest amendment of any given rule.

EFFECTIVE AND EXPIRATION DATES

    Each volume of the Code contains amendments published in the Federal 
Register since the last revision of that volume of the Code. Source 
citations for the regulations are referred to by volume number and page 
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Code a note has been inserted to reflect the future effective date. In 
those instances where a regulation published in the Federal Register 
states a date certain for expiration, an appropriate note will be 
inserted following the text.

OMB CONTROL NUMBERS

    The Paperwork Reduction Act of 1980 (Pub. L. 96-511) requires 
Federal agencies to display an OMB control number with their information 
collection request.

[[Page vi]]

Many agencies have begun publishing numerous OMB control numbers as 
amendments to existing regulations in the CFR. These OMB numbers are 
placed as close as possible to the applicable recordkeeping or reporting 
requirements.

PAST PROVISIONS OF THE CODE

    Provisions of the Code that are no longer in force and effect as of 
the revision date stated on the cover of each volume are not carried. 
Code users may find the text of provisions in effect on any given date 
in the past by using the appropriate List of CFR Sections Affected 
(LSA). For the convenience of the reader, a ``List of CFR Sections 
Affected'' is published at the end of each CFR volume. For changes to 
the Code prior to the LSA listings at the end of the volume, consult 
previous annual editions of the LSA. For changes to the Code prior to 
2001, consult the List of CFR Sections Affected compilations, published 
for 1949-1963, 1964-1972, 1973-1985, and 1986-2000.

``[RESERVED]'' TERMINOLOGY

    The term ``[Reserved]'' is used as a place holder within the Code of 
Federal Regulations. An agency may add regulatory information at a 
``[Reserved]'' location at any time. Occasionally ``[Reserved]'' is used 
editorially to indicate that a portion of the CFR was left vacant and 
not accidentally dropped due to a printing or computer error.

INCORPORATION BY REFERENCE

    What is incorporation by reference? Incorporation by reference was 
established by statute and allows Federal agencies to meet the 
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to materials already published elsewhere. For an incorporation to be 
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This material, like any other properly issued regulation, has the force 
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    What is a proper incorporation by reference? The Director of the 
Federal Register will approve an incorporation by reference only when 
the requirements of 1 CFR part 51 are met. Some of the elements on which 
approval is based are:
    (a) The incorporation will substantially reduce the volume of 
material published in the Federal Register.
    (b) The matter incorporated is in fact available to the extent 
necessary to afford fairness and uniformity in the administrative 
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    (c) The incorporating document is drafted and submitted for 
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CFR INDEXES AND TABULAR GUIDES

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and Finding Aids. This volume contains the Parallel Table of Authorities 
and Rules. A list of CFR titles, chapters, subchapters, and parts and an 
alphabetical list of agencies publishing in the CFR are also included in 
this volume.

[[Page vii]]

    An index to the text of ``Title 3--The President'' is carried within 
that volume.
    The Federal Register Index is issued monthly in cumulative form. 
This index is based on a consolidation of the ``Contents'' entries in 
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    A List of CFR Sections Affected (LSA) is published monthly, keyed to 
the revision dates of the 50 CFR titles.

REPUBLICATION OF MATERIAL

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in the Code of Federal Regulations.

INQUIRIES

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    Amy P. Bunk,
    Acting Director,
    Office of the Federal Register.
    January 1, 2015.







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                               THIS TITLE

    Title 14--Aeronautics and Space is composed of five volumes. The 
parts in these volumes are arranged in the following order: Parts 1-59, 
60-109, 110-199, 200-1199, and part 1200-End. The first three volumes 
containing parts 1-199 are comprised of chapter I--Federal Aviation 
Administration, Department of Transportation (DOT). The fourth volume 
containing parts 200-1199 is comprised of chapter II--Office of the 
Secretary, DOT (Aviation Proceedings) and chapter III--Commercial Space 
Transportation, Federal Aviation Administration, DOT. The fifth volume 
containing part 1200-End is comprised of chapter V--National Aeronautics 
and Space Administration and chapter VI--Air Transportation System 
Stabilization. The contents of these volumes represent all current 
regulations codified under this title of the CFR as of January 1, 2015.

    For this volume, Susannah C. Hurley was Chief Editor. The Code of 
Federal Regulations publication program is under the direction of John 
Hyrum Martinez, assisted by Stephen J. Frattini.

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                     TITLE 14--AERONAUTICS AND SPACE




                   (This book contains parts 1 to 59)

  --------------------------------------------------------------------
                                                                    Part

chapter i--Federal Aviation Administration, Department of 
  Transportation............................................           1

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CHAPTER I--FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION




  --------------------------------------------------------------------

                        SUBCHAPTER A--DEFINITIONS
Part                                                                Page
1               Definitions and abbreviations...............           5
3               General requirements........................          20
                     SUBCHAPTER B--PROCEDURAL RULES
11              General rulemaking procedures...............          22
13              Investigative and enforcement procedures....          33
14              Rules implementing the Equal Access to 
                    Justice Act of 1980.....................          78
15              Administrative claims under Federal Tort 
                    Claims Act..............................          84
16              Rules of practice for Federally-assisted 
                    airport enforcement proceedings.........          89
17              Procedures for protests and contract 
                    disputes................................         110
                         SUBCHAPTER C--AIRCRAFT
21              Certification procedures for products and 
                    parts...................................         127
23              Airworthiness standards: Normal, utility, 
                    acrobatic, and commuter category 
                    airplanes...............................         170
25              Airworthiness standards: Transport category 
                    airplanes...............................         360
26              Continued airworthiness and safety 
                    improvements for transport category 
                    airplanes...............................         642
27              Airworthiness standards: Normal category 
                    rotorcraft..............................         657
29              Airworthiness standards: Transport category 
                    rotorcraft..............................         743
31              Airworthiness standards: Manned free 
                    balloons................................         862
33              Airworthiness standards: Aircraft engines...         869
34              Fuel venting and exhaust emission 
                    requirements for turbine engine powered 
                    airplanes...............................         915
35              Airworthiness standards: Propellers.........         927

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36              Noise standards: Aircraft type and 
                    airworthiness certification.............         935
39              Airworthiness directives....................        1013
43              Maintenance, preventive maintenance, 
                    rebuilding, and alteration..............        1015
45              Identification and registration marking.....        1031
47              Aircraft registration.......................        1037
49              Recording of aircraft titles and security 
                    documents...............................        1051
50-59           [Reserved]

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                        SUBCHAPTER A_DEFINITIONS





PART 1_DEFINITIONS AND ABBREVIATIONS--Table of Contents



Sec.
1.1 General definitions.
1.2 Abbreviations and symbols.
1.3 Rules of construction.

    Authority: 49 U.S.C. 106(g), 40113, 44701.



Sec. 1.1  General definitions.

    As used in Subchapters A through K of this chapter, unless the 
context requires otherwise:
    Administrator means the Federal Aviation Administrator or any person 
to whom he has delegated his authority in the matter concerned.
    Aerodynamic coefficients means non-dimensional coefficients for 
aerodynamic forces and moments.
    Air carrier means a person who undertakes directly by lease, or 
other arrangement, to engage in air transportation.
    Air commerce means interstate, overseas, or foreign air commerce or 
the transportation of mail by aircraft or any operation or navigation of 
aircraft within the limits of any Federal airway or any operation or 
navigation of aircraft which directly affects, or which may endanger 
safety in, interstate, overseas, or foreign air commerce.
    Aircraft means a device that is used or intended to be used for 
flight in the air.
    Aircraft engine means an engine that is used or intended to be used 
for propelling aircraft. It includes turbosuperchargers, appurtenances, 
and accessories necessary for its functioning, but does not include 
propellers.
    Airframe means the fuselage, booms, nacelles, cowlings, fairings, 
airfoil surfaces (including rotors but excluding propellers and rotating 
airfoils of engines), and landing gear of an aircraft and their 
accessories and controls.
    Airplane means an engine-driven fixed-wing aircraft heavier than 
air, that is supported in flight by the dynamic reaction of the air 
against its wings.
    Airport means an area of land or water that is used or intended to 
be used for the landing and takeoff of aircraft, and includes its 
buildings and facilities, if any.
    Airship means an engine-driven lighter-than-air aircraft that can be 
steered.
    Air traffic means aircraft operating in the air or on an airport 
surface, exclusive of loading ramps and parking areas.
    Air traffic clearance means an authorization by air traffic control, 
for the purpose of preventing collision between known aircraft, for an 
aircraft to proceed under specified traffic conditions within controlled 
airspace.
    Air traffic control means a service operated by appropriate 
authority to promote the safe, orderly, and expeditious flow of air 
traffic.
    Air Traffic Service (ATS) route is a specified route designated for 
channeling the flow of traffic as necessary for the provision of air 
traffic services. The term ``ATS route'' refers to a variety of airways, 
including jet routes, area navigation (RNAV) routes, and arrival and 
departure routes. An ATS route is defined by route specifications, which 
may include:
    (1) An ATS route designator;
    (2) The path to or from significant points;
    (3) Distance between significant points;
    (4) Reporting requirements; and
    (5) The lowest safe altitude determined by the appropriate 
authority.
    Air transportation means interstate, overseas, or foreign air 
transportation or the transportation of mail by aircraft.
    Alert Area. An alert area is established to inform pilots of a 
specific area wherein a high volume of pilot training or an unusual type 
of aeronautical activity is conducted.
    Alternate airport means an airport at which an aircraft may land if 
a landing at the intended airport becomes inadvisable.
    Altitude engine means a reciprocating aircraft engine having a rated 
takeoff power that is producible from sea level to an established higher 
altitude.
    Amateur rocket means an unmanned rocket that:

[[Page 6]]

    (1) Is propelled by a motor or motors having a combined total 
impulse of 889,600 Newton-seconds (200,000 pound-seconds) or less; and
    (2) Cannot reach an altitude greater than 150 kilometers (93.2 
statute miles) above the earth's surface.
    Appliance means any instrument, mechanism, equipment, part, 
apparatus, appurtenance, or accessory, including communications 
equipment, that is used or intended to be used in operating or 
controlling an aircraft in flight, is installed in or attached to the 
aircraft, and is not part of an airframe, engine, or propeller.
    Approved, unless used with reference to another person, means 
approved by the FAA or any person to whom the FAA has delegated its 
authority in the matter concerned, or approved under the provisions of a 
bilateral agreement between the United States and a foreign country or 
jurisdiction.
    Area navigation (RNAV) is a method of navigation that permits 
aircraft operations on any desired flight path.
    Area navigation (RNAV) route is an ATS route based on RNAV that can 
be used by suitably equipped aircraft.
    Armed Forces means the Army, Navy, Air Force, Marine Corps, and 
Coast Guard, including their regular and reserve components and members 
serving without component status.
    Autorotation means a rotorcraft flight condition in which the 
lifting rotor is driven entirely by action of the air when the 
rotorcraft is in motion.
    Auxiliary rotor means a rotor that serves either to counteract the 
effect of the main rotor torque on a rotorcraft or to maneuver the 
rotorcraft about one or more of its three principal axes.
    Balloon means a lighter-than-air aircraft that is not engine driven, 
and that sustains flight through the use of either gas buoyancy or an 
airborne heater.
    Brake horsepower means the power delivered at the propeller shaft 
(main drive or main output) of an aircraft engine.
    Calibrated airspeed means the indicated airspeed of an aircraft, 
corrected for position and instrument error. Calibrated airspeed is 
equal to true airspeed in standard atmosphere at sea level.
    Canard means the forward wing of a canard configuration and may be a 
fixed, movable, or variable geometry surface, with or without control 
surfaces.
    Canard configuration means a configuration in which the span of the 
forward wing is substantially less than that of the main wing.
    Category:
    (1) As used with respect to the certification, ratings, privileges, 
and limitations of airmen, means a broad classification of aircraft. 
Examples include: airplane; rotorcraft; glider; and lighter-than-air; 
and
    (2) As used with respect to the certification of aircraft, means a 
grouping of aircraft based upon intended use or operating limitations. 
Examples include: transport, normal, utility, acrobatic, limited, 
restricted, and provisional.
    Category A, with respect to transport category rotorcraft, means 
multiengine rotorcraft designed with engine and system isolation 
features specified in Part 29 and utilizing scheduled takeoff and 
landing operations under a critical engine failure concept which assures 
adequate designated surface area and adequate performance capability for 
continued safe flight in the event of engine failure.
    Category B, with respect to transport category rotorcraft, means 
single-engine or multiengine rotorcraft which do not fully meet all 
Category A standards. Category B rotorcraft have no guaranteed stay-up 
ability in the event of engine failure and unscheduled landing is 
assumed.
    Category II operations, with respect to the operation of aircraft, 
means a straight-in ILS approach to the runway of an airport under a 
Category II ILS instrument approach procedure issued by the 
Administrator or other appropriate authority.
    Category III operations, with respect to the operation of aircraft, 
means an ILS approach to, and landing on, the runway of an airport using 
a Category III ILS instrument approach procedure issued by the 
Administrator or other appropriate authority.
    Ceiling means the height above the earth's surface of the lowest 
layer of clouds or obscuring phenomena that is

[[Page 7]]

reported as ``broken'', ``overcast'', or ``obscuration'', and not 
classified as ``thin'' or ``partial''.
    Civil aircraft means aircraft other than public aircraft.
    Class:
    (1) As used with respect to the certification, ratings, privileges, 
and limitations of airmen, means a classification of aircraft within a 
category having similar operating characteristics. Examples include: 
single engine; multiengine; land; water; gyroplane; helicopter; airship; 
and free balloon; and
    (2) As used with respect to the certification of aircraft, means a 
broad grouping of aircraft having similar characteristics of propulsion, 
flight, or landing. Examples include: airplane; rotorcraft; glider; 
balloon; landplane; and seaplane.
    Clearway means:
    (1) For turbine engine powered airplanes certificated after August 
29, 1959, an area beyond the runway, not less than 500 feet wide, 
centrally located about the extended centerline of the runway, and under 
the control of the airport authorities. The clearway is expressed in 
terms of a clearway plane, extending from the end of the runway with an 
upward slope not exceeding 1.25 percent, above which no object nor any 
terrain protrudes. However, threshold lights may protrude above the 
plane if their height above the end of the runway is 26 inches or less 
and if they are located to each side of the runway.
    (2) For turbine engine powered airplanes certificated after 
September 30, 1958, but before August 30, 1959, an area beyond the 
takeoff runway extending no less than 300 feet on either side of the 
extended centerline of the runway, at an elevation no higher than the 
elevation of the end of the runway, clear of all fixed obstacles, and 
under the control of the airport authorities.
    Climbout speed, with respect to rotorcraft, means a referenced 
airspeed which results in a flight path clear of the height-velocity 
envelope during initial climbout.
    Commercial operator means a person who, for compensation or hire, 
engages in the carriage by aircraft in air commerce of persons or 
property, other than as an air carrier or foreign air carrier or under 
the authority of Part 375 of this title. Where it is doubtful that an 
operation is for ``compensation or hire'', the test applied is whether 
the carriage by air is merely incidental to the person's other business 
or is, in itself, a major enterprise for profit.
    Configuration, Maintenance, and Procedures (CMP) document means a 
document approved by the FAA that contains minimum configuration, 
operating, and maintenance requirements, hardware life-limits, and 
Master Minimum Equipment List (MMEL) constraints necessary for an 
airplane-engine combination to meet ETOPS type design approval 
requirements.
    Consensus standard means, for the purpose of certificating light-
sport aircraft, an industry-developed consensus standard that applies to 
aircraft design, production, and airworthiness. It includes, but is not 
limited to, standards for aircraft design and performance, required 
equipment, manufacturer quality assurance systems, production acceptance 
test procedures, operating instructions, maintenance and inspection 
procedures, identification and recording of major repairs and major 
alterations, and continued airworthiness.
    Controlled airspace means an airspace of defined dimensions within 
which air traffic control service is provided to IFR flights and to VFR 
flights in accordance with the airspace classification.

    Note: Controlled airspace is a generic term that covers Class A, 
Class B, Class C, Class D, and Class E airspace.

    Controlled Firing Area. A controlled firing area is established to 
contain activities, which if not conducted in a controlled environment, 
would be hazardous to nonparticipating aircraft.
    Crewmember means a person assigned to perform duty in an aircraft 
during flight time.
    Critical altitude means the maximum altitude at which, in standard 
atmosphere, it is possible to maintain, at a specified rotational speed, 
a specified power or a specified manifold pressure. Unless otherwise 
stated, the critical altitude is the maximum altitude at which it is 
possible to maintain, at the maximum continuous rotational speed, one of 
the following:

[[Page 8]]

    (1) The maximum continuous power, in the case of engines for which 
this power rating is the same at sea level and at the rated altitude.
    (2) The maximum continuous rated manifold pressure, in the case of 
engines, the maximum continuous power of which is governed by a constant 
manifold pressure.
    Critical engine means the engine whose failure would most adversely 
affect the performance or handling qualities of an aircraft.
    Decision altitude (DA) is a specified altitude in an instrument 
approach procedure at which the pilot must decide whether to initiate an 
immediate missed approach if the pilot does not see the required visual 
reference, or to continue the approach. Decision altitude is expressed 
in feet above mean sea level.
    Decision height (DH) is a specified height above the ground in an 
instrument approach procedure at which the pilot must decide whether to 
initiate an immediate missed approach if the pilot does not see the 
required visual reference, or to continue the approach. Decision height 
is expressed in feet above ground level.
    Early ETOPS means ETOPS type design approval obtained without 
gaining non-ETOPS service experience on the candidate airplane-engine 
combination certified for ETOPS.
    Enhanced flight visibility (EFV) means the average forward 
horizontal distance, from the cockpit of an aircraft in flight, at which 
prominent topographical objects may be clearly distinguished and 
identified by day or night by a pilot using an enhanced flight vision 
system.
    Enhanced flight vision system (EFVS) means an electronic means to 
provide a display of the forward external scene topography (the natural 
or manmade features of a place or region especially in a way to show 
their relative positions and elevation) through the use of imaging 
sensors, such as a forward looking infrared, millimeter wave radiometry, 
millimeter wave radar, low light level image intensifying.
    Equivalent airspeed means the calibrated airspeed of an aircraft 
corrected for adiabatic compressible flow for the particular altitude. 
Equivalent airspeed is equal to calibrated airspeed in standard 
atmosphere at sea level.
    ETOPS Significant System means an airplane system, including the 
propulsion system, the failure or malfunctioning of which could 
adversely affect the safety of an ETOPS flight, or the continued safe 
flight and landing of an airplane during an ETOPS diversion. Each ETOPS 
significant system is either an ETOPS group 1 significant system or an 
ETOPS group 2 significant system.
    (1) An ETOPS group 1 Significant System--
    (i) Has fail-safe characteristics directly linked to the degree of 
redundancy provided by the number of engines on the airplane.
    (ii) Is a system, the failure or malfunction of which could result 
in an IFSD, loss of thrust control, or other power loss.
    (iii) Contributes significantly to the safety of an ETOPS diversion 
by providing additional redundancy for any system power source lost as a 
result of an inoperative engine.
    (iv) Is essential for prolonged operation of an airplane at engine 
inoperative altitudes.
    (2) An ETOPS group 2 significant system is an ETOPS significant 
system that is not an ETOPS group 1 significant system.
    Extended Operations (ETOPS) means an airplane flight operation, 
other than an all-cargo operation in an airplane with more than two 
engines, during which a portion of the flight is conducted beyond a time 
threshold identified in part 121 or part 135 of this chapter that is 
determined using an approved one-engine-inoperative cruise speed under 
standard atmospheric conditions in still air.
    Extended over-water operation means--
    (1) With respect to aircraft other than helicopters, an operation 
over water at a horizontal distance of more than 50 nautical miles from 
the nearest shoreline; and
    (2) With respect to helicopters, an operation over water at a 
horizontal distance of more than 50 nautical miles from the nearest 
shoreline and more than 50 nautical miles from an off-shore heliport 
structure.

[[Page 9]]

    External load means a load that is carried, or extends, outside of 
the aircraft fuselage.
    External-load attaching means means the structural components used 
to attach an external load to an aircraft, including external-load 
containers, the backup structure at the attachment points, and any 
quick-release device used to jettison the external load.
    Final approach fix (FAF) defines the beginning of the final approach 
segment and the point where final segment descent may begin.
    Final takeoff speed means the speed of the airplane that exists at 
the end of the takeoff path in the en route configuration with one 
engine inoperative.
    Fireproof--
    (1) With respect to materials and parts used to confine fire in a 
designated fire zone, means the capacity to withstand at least as well 
as steel in dimensions appropriate for the purpose for which they are 
used, the heat produced when there is a severe fire of extended duration 
in that zone; and
    (2) With respect to other materials and parts, means the capacity to 
withstand the heat associated with fire at least as well as steel in 
dimensions appropriate for the purpose for which they are used.
    Fire resistant--
    (1) With respect to sheet or structural members means the capacity 
to withstand the heat associated with fire at least as well as aluminum 
alloy in dimensions appropriate for the purpose for which they are used; 
and
    (2) With respect to fluid-carrying lines, fluid system parts, 
wiring, air ducts, fittings, and powerplant controls, means the capacity 
to perform the intended functions under the heat and other conditions 
likely to occur when there is a fire at the place concerned.
    Flame resistant means not susceptible to combustion to the point of 
propagating a flame, beyond safe limits, after the ignition source is 
removed.
    Flammable, with respect to a fluid or gas, means susceptible to 
igniting readily or to exploding.
    Flap extended speed means the highest speed permissible with wing 
flaps in a prescribed extended position.
    Flash resistant means not susceptible to burning violently when 
ignited.
    Flightcrew member means a pilot, flight engineer, or flight 
navigator assigned to duty in an aircraft during flight time.
    Flight level means a level of constant atmospheric pressure related 
to a reference datum of 29.92 inches of mercury. Each is stated in three 
digits that represent hundreds of feet. For example, flight level 250 
represents a barometric altimeter indication of 25,000 feet; flight 
level 255, an indication of 25,500 feet.
    Flight plan means specified information, relating to the intended 
flight of an aircraft, that is filed orally or in writing with air 
traffic control.
    Flight simulation training device (FSTD) means a flight simulator or 
a flight training device.
    Flight time means:
    (1) Pilot time that commences when an aircraft moves under its own 
power for the purpose of flight and ends when the aircraft comes to rest 
after landing; or
    (2) For a glider without self-launch capability, pilot time that 
commences when the glider is towed for the purpose of flight and ends 
when the glider comes to rest after landing.
    Flight training device (FTD) means a replica of aircraft 
instruments, equipment, panels, and controls in an open flight deck area 
or an enclosed aircraft cockpit replica. It includes the equipment and 
computer programs necessary to represent aircraft (or set of aircraft) 
operations in ground and flight conditions having the full range of 
capabilities of the systems installed in the device as described in part 
60 of this chapter and the qualification performance standard (QPS) for 
a specific FTD qualification level.
    Flight visibility means the average forward horizontal distance, 
from the cockpit of an aircraft in flight, at which prominent unlighted 
objects may be seen and identified by day and prominent lighted objects 
may be seen and identified by night.
    Foreign air carrier means any person other than a citizen of the 
United States, who undertakes directly, by lease or other arrangement, 
to engage in air transportation.

[[Page 10]]

    Foreign air commerce means the carriage by aircraft of persons or 
property for compensation or hire, or the carriage of mail by aircraft, 
or the operation or navigation of aircraft in the conduct or furtherance 
of a business or vocation, in commerce between a place in the United 
States and any place outside thereof; whether such commerce moves wholly 
by aircraft or partly by aircraft and partly by other forms of 
transportation.
    Foreign air transportation means the carriage by aircraft of persons 
or property as a common carrier for compensation or hire, or the 
carriage of mail by aircraft, in commerce between a place in the United 
States and any place outside of the United States, whether that commerce 
moves wholly by aircraft or partly by aircraft and partly by other forms 
of transportation.
    Forward wing means a forward lifting surface of a canard 
configuration or tandem-wing configuration airplane. The surface may be 
a fixed, movable, or variable geometry surface, with or without control 
surfaces.
    Full flight simulator (FFS) means a replica of a specific type; or 
make, model, and series aircraft cockpit. It includes the assemblage of 
equipment and computer programs necessary to represent aircraft 
operations in ground and flight conditions, a visual system providing an 
out-of-the-cockpit view, a system that provides cues at least equivalent 
to those of a three-degree-of-freedom motion system, and has the full 
range of capabilities of the systems installed in the device as 
described in part 60 of this chapter and the qualification performance 
standards (QPS) for a specific FFS qualification level.
    Glider means a heavier-than-air aircraft, that is supported in 
flight by the dynamic reaction of the air against its lifting surfaces 
and whose free flight does not depend principally on an engine.
    Ground visibility means prevailing horizontal visibility near the 
earth's surface as reported by the United States National Weather 
Service or an accredited observer.
    Go-around power or thrust setting means the maximum allowable in-
flight power or thrust setting identified in the performance data.
    Gyrodyne means a rotorcraft whose rotors are normally engine-driven 
for takeoff, hovering, and landing, and for forward flight through part 
of its speed range, and whose means of propulsion, consisting usually of 
conventional propellers, is independent of the rotor system.
    Gyroplane means a rotorcraft whose rotors are not engine-driven, 
except for initial starting, but are made to rotate by action of the air 
when the rotorcraft is moving; and whose means of propulsion, consisting 
usually of conventional propellers, is independent of the rotor system.
    Helicopter means a rotorcraft that, for its horizontal motion, 
depends principally on its engine-driven rotors.
    Heliport means an area of land, water, or structure used or intended 
to be used for the landing and takeoff of helicopters.
    Idle thrust means the jet thrust obtained with the engine power 
control level set at the stop for the least thrust position at which it 
can be placed.
    IFR conditions means weather conditions below the minimum for flight 
under visual flight rules.
    IFR over-the-top, with respect to the operation of aircraft, means 
the operation of an aircraft over-the-top on an IFR flight plan when 
cleared by air traffic control to maintain ``VFR conditions'' or ``VFR 
conditions on top''.
    Indicated airspeed means the speed of an aircraft as shown on its 
pitot static airspeed indicator calibrated to reflect standard 
atmosphere adiabatic compressible flow at sea level uncorrected for 
airspeed system errors.
    In-flight shutdown (IFSD) means, for ETOPS only, when an engine 
ceases to function (when the airplane is airborne) and is shutdown, 
whether self induced, flightcrew initiated or caused by an external 
influence. The FAA considers IFSD for all causes: for example, flameout, 
internal failure, flightcrew initiated shutdown, foreign object 
ingestion, icing, inability to obtain or control desired thrust or 
power, and cycling of the start control, however briefly, even if the 
engine operates normally for the remainder of the flight. This 
definition excludes the airborne

[[Page 11]]

cessation of the functioning of an engine when immediately followed by 
an automatic engine relight and when an engine does not achieve desired 
thrust or power but is not shutdown.
    Instrument means a device using an internal mechanism to show 
visually or aurally the attitude, altitude, or operation of an aircraft 
or aircraft part. It includes electronic devices for automatically 
controlling an aircraft in flight.
    Instrument approach procedure (IAP) is a series of predetermined 
maneuvers by reference to flight instruments with specified protection 
from obstacles and assurance of navigation signal reception capability. 
It begins from the initial approach fix, or where applicable, from the 
beginning of a defined arrival route to a point:
    (1) From which a landing can be completed; or
    (2) If a landing is not completed, to a position at which holding or 
en route obstacle clearance criteria apply.
    Interstate air commerce means the carriage by aircraft of persons or 
property for compensation or hire, or the carriage of mail by aircraft, 
or the operation or navigation of aircraft in the conduct or furtherance 
of a business or vocation, in commerce between a place in any State of 
the United States, or the District of Columbia, and a place in any other 
State of the United States, or the District of Columbia; or between 
places in the same State of the United States through the airspace over 
any place outside thereof; or between places in the same territory or 
possession of the United States, or the District of Columbia.
    Interstate air transportation means the carriage by aircraft of 
persons or property as a common carrier for compensation or hire, or the 
carriage of mail by aircraft in commerce:
    (1) Between a place in a State or the District of Columbia and 
another place in another State or the District of Columbia;
    (2) Between places in the same State through the airspace over any 
place outside that State; or
    (3) Between places in the same possession of the United States;

Whether that commerce moves wholly by aircraft of partly by aircraft and 
partly by other forms of transportation.
    Intrastate air transportation means the carriage of persons or 
property as a common carrier for compensation or hire, by turbojet-
powered aircraft capable of carrying thirty or more persons, wholly 
within the same State of the United States.
    Kite means a framework, covered with paper, cloth, metal, or other 
material, intended to be flown at the end of a rope or cable, and having 
as its only support the force of the wind moving past its surfaces.
    Landing gear extended speed means the maximum speed at which an 
aircraft can be safely flown with the landing gear extended.
    Landing gear operating speed means the maximum speed at which the 
landing gear can be safely extended or retracted.
    Large aircraft means aircraft of more than 12,500 pounds, maximum 
certificated takeoff weight.
    Light-sport aircraft means an aircraft, other than a helicopter or 
powered-lift that, since its original certification, has continued to 
meet the following:
    (1) A maximum takeoff weight of not more than--
    (i) 1,320 pounds (600 kilograms) for aircraft not intended for 
operation on water; or
    (ii) 1,430 pounds (650 kilograms) for an aircraft intended for 
operation on water.
    (2) A maximum airspeed in level flight with maximum continuous power 
(VH) of not more than 120 knots CAS under standard 
atmospheric conditions at sea level.
    (3) A maximum never-exceed speed (VNE) of not more than 
120 knots CAS for a glider.
    (4) A maximum stalling speed or minimum steady flight speed without 
the use of lift-enhancing devices (VS1) of not more than 45 
knots CAS at the aircraft's maximum certificated takeoff weight and most 
critical center of gravity.
    (5) A maximum seating capacity of no more than two persons, 
including the pilot.
    (6) A single, reciprocating engine, if powered.

[[Page 12]]

    (7) A fixed or ground-adjustable propeller if a powered aircraft 
other than a powered glider.
    (8) A fixed or feathering propeller system if a powered glider.
    (9) A fixed-pitch, semi-rigid, teetering, two-blade rotor system, if 
a gyroplane.
    (10) A nonpressurized cabin, if equipped with a cabin.
    (11) Fixed landing gear, except for an aircraft intended for 
operation on water or a glider.
    (12) Fixed or retractable landing gear, or a hull, for an aircraft 
intended for operation on water.
    (13) Fixed or retractable landing gear for a glider.
    Lighter-than-air aircraft means aircraft that can rise and remain 
suspended by using contained gas weighing less than the air that is 
displaced by the gas.
    Load factor means the ratio of a specified load to the total weight 
of the aircraft. The specified load is expressed in terms of any of the 
following: aerodynamic forces, inertia forces, or ground or water 
reactions.
    Long-range communication system (LRCS). A system that uses satellite 
relay, data link, high frequency, or another approved communication 
system which extends beyond line of sight.
    Long-range navigation system (LRNS). An electronic navigation unit 
that is approved for use under instrument flight rules as a primary 
means of navigation, and has at least one source of navigational input, 
such as inertial navigation system, global positioning system, Omega/
very low frequency, or Loran C.
    Mach number means the ratio of true airspeed to the speed of sound.
    Main rotor means the rotor that supplies the principal lift to a 
rotorcraft.
    Maintenance means inspection, overhaul, repair, preservation, and 
the replacement of parts, but excludes preventive maintenance.
    Major alteration means an alteration not listed in the aircraft, 
aircraft engine, or propeller specifications--
    (1) That might appreciably affect weight, balance, structural 
strength, performance, powerplant operation, flight characteristics, or 
other qualities affecting airworthiness; or
    (2) That is not done according to accepted practices or cannot be 
done by elementary operations.
    Major repair means a repair:
    (1) That, if improperly done, might appreciably affect weight, 
balance, structural strength, performance, powerplant operation, flight 
characteristics, or other qualities affecting airworthiness; or
    (2) That is not done according to accepted practices or cannot be 
done by elementary operations.
    Manifold pressure means absolute pressure as measured at the 
appropriate point in the induction system and usually expressed in 
inches of mercury.
    Maximum engine overtorque, as it applies to turbopropeller and 
turboshaft engines incorporating free power turbines for all ratings 
except one engine inoperative (OEI) ratings of two minutes or less, 
means the maximum torque of the free power turbine rotor assembly, the 
inadvertent occurrence of which, for periods of up to 20 seconds, will 
not require rejection of the engine from service, or any maintenance 
action other than to correct the cause.Maximum speed for stability 
characteristics, VFC/MFC means a speed that may 
not be less than a speed midway between maximum operating limit speed 
(VMO/MMO) and demonstrated flight diving speed 
(VDF/MDF), except that, for altitudes where the 
Mach number is the limiting factor, MFC need not exceed the 
Mach number at which effective speed warning occurs.
    Medical certificate means acceptable evidence of physical fitness on 
a form prescribed by the Administrator.
    Military operations area. A military operations area (MOA) is 
airspace established outside Class A airspace to separate or segregate 
certain nonhazardous military activities from IFR Traffic and to 
identify for VFR traffic where theses activities are conducted.
    VA means design maneuvering speed.
    VB means design speed for maximum gust intensity.
    VC means design cruising speed.
    VD means design diving speed.
    VDF/MDF means demonstrated flight diving speed.

[[Page 13]]

    VEF means the speed at which the critical engine is assumed to fail 
during takeoff.
    VF means design flap speed.
    VFC/MFC means maximum speed for stability characteristics.
    VFE means maximum flap extended speed.
    VH means maximum speed in level flight with maximum continuous 
power.
    VLE means maximum landing gear extended speed.
    VLO means maximum landing gear operating speed.
    VLOF means lift-off speed.
    VMC means minimum control speed with the critical engine 
inoperative.
    VMO/MMO means maximum operating limit speed.
    VMU means minimum unstick speed.
    VNE means never-exceed speed.
    VNO means maximum structural cruising speed.
    VR means rotation speed.
    VS means the stalling speed or the minimum steady flight speed at 
which the airplane is controllable.
    Minimum descent altitude (MDA) is the lowest altitude specified in 
an instrument approach procedure, expressed in feet above mean sea 
level, to which descent is authorized on final approach or during 
circle-to-land maneuvering until the pilot sees the required visual 
references for the heliport or runway of intended landing.
    Minor alteration means an alteration other than a major alteration.
    Minor repair means a repair other than a major repair.
    National defense airspace means airspace established by a regulation 
prescribed, or an order issued under, 49 U.S.C. 40103(b)(3).
    Navigable airspace means airspace at and above the minimum flight 
altitudes prescribed by or under this chapter, including airspace needed 
for safe takeoff and landing.
    Night means the time between the end of evening civil twilight and 
the beginning of morning civil twilight, as published in the Air 
Almanac, converted to local time.
    Nonprecision approach procedure means a standard instrument approach 
procedure in which no electronic glide slope is provided.
    Operate, with respect to aircraft, means use, cause to use or 
authorize to use aircraft, for the purpose (except as provided in Sec. 
91.13 of this chapter) of air navigation including the piloting of 
aircraft, with or without the right of legal control (as owner, lessee, 
or otherwise).
    Operational control, with respect to a flight, means the exercise of 
authority over initiating, conducting or terminating a flight.
    Overseas air commerce means the carriage by aircraft of persons or 
property for compensation or hire, or the carriage of mail by aircraft, 
or the operation or navigation of aircraft in the conduct or furtherance 
of a business or vocation, in commerce between a place in any State of 
the United States, or the District of Columbia, and any place in a 
territory or possession of the United States; or between a place in a 
territory or possession of the United States, and a place in any other 
territory or possession of the United States.
    Overseas air transportation means the carriage by aircraft of 
persons or property as a common carrier for compensation or hire, or the 
carriage of mail by aircraft, in commerce:
    (1) Between a place in a State or the District of Columbia and a 
place in a possession of the United States; or
    (2) Between a place in a possession of the United States and a place 
in another possession of the United States; whether that commerce moves 
wholly by aircraft or partly by aircraft and partly by other forms of 
transportation.
    Over-the-top means above the layer of clouds or other obscuring 
phenomena forming the ceiling.
    Parachute means a device used or intended to be used to retard the 
fall of a body or object through the air.
    Person means an individual, firm, partnership, corporation, company, 
association, joint-stock association, or governmental entity. It 
includes a trustee, receiver, assignee, or similar representative of any 
of them.
    Pilotage means navigation by visual reference to landmarks.
    Pilot in command means the person who:

[[Page 14]]

    (1) Has final authority and responsibility for the operation and 
safety of the flight;
    (2) Has been designated as pilot in command before or during the 
flight; and
    (3) Holds the appropriate category, class, and type rating, if 
appropriate, for the conduct of the flight.
    Pitch setting means the propeller blade setting as determined by the 
blade angle measured in a manner, and at a radius, specified by the 
instruction manual for the propeller.
    Positive control means control of all air traffic, within designated 
airspace, by air traffic control.
    Powered parachute means a powered aircraft comprised of a flexible 
or semi-rigid wing connected to a fuselage so that the wing is not in 
position for flight until the aircraft is in motion. The fuselage of a 
powered parachute contains the aircraft engine, a seat for each occupant 
and is attached to the aircraft's landing gear.
    Powered-lift means a heavier-than-air aircraft capable of vertical 
takeoff, vertical landing, and low speed flight that depends principally 
on engine-driven lift devices or engine thrust for lift during these 
flight regimes and on nonrotating airfoil(s) for lift during horizontal 
flight.
    Precision approach procedure means a standard instrument approach 
procedure in which an electronic glide slope is provided, such as ILS 
and PAR.
    Preventive maintenance means simple or minor preservation operations 
and the replacement of small standard parts not involving complex 
assembly operations.
    Prohibited area. A prohibited area is airspace designated under part 
73 within which no person may operate an aircraft without the permission 
of the using agency.
    Propeller means a device for propelling an aircraft that has blades 
on an engine-driven shaft and that, when rotated, produces by its action 
on the air, a thrust approximately perpendicular to its plane of 
rotation. It includes control components normally supplied by its 
manufacturer, but does not include main and auxiliary rotors or rotating 
airfoils of engines.
    Public aircraft means any of the following aircraft when not being 
used for a commercial purpose or to carry an individual other than a 
crewmember or qualified non-crewmenber:
    (1) An aircraft used only for the United States Government; an 
aircraft owned by the Government and operated by any person for purposes 
related to crew training, equipment development, or demonstration; an 
aircraft owned and operated by the government of a State, the District 
of Columbia, or a territory or possession of the United States or a 
political subdivision of one of these governments; or an aircraft 
exclusively leased for at least 90 continuous days by the government of 
a State, the District of Columbia, or a territory or possession of the 
United States or a political subdivision of one of these governments.
    (i) For the sole purpose of determining public aircraft status, 
commercial purposes means the transportation of persons or property for 
compensation or hire, but does not include the operation of an aircraft 
by the armed forces for reimbursement when that reimbursement is 
required by any Federal statute, regulation, or directive, in effect on 
November 1, 1999, or by one government on behalf of another government 
under a cost reimbursement agreement if the government on whose behalf 
the operation is conducted certifies to the Administrator of the Federal 
Aviation Administration that the operation is necessary to respond to a 
significant and imminent threat to life or property (including natural 
resources) and that no service by a private operator is reasonably 
available to meet the threat.
    (ii) For the sole purpose of determining public aircraft status, 
governmental function means an activity undertaken by a government, such 
as national defense, intelligence missions, firefighting, search and 
rescue, law enforcement (including transport of prisoners, detainees, 
and illegal aliens), aeronautical research, or biological or geological 
resource management.
    (iii) For the sole purpose of determining public aircraft status, 
qualified non-crewmember means an individual, other than a member of the 
crew, aboard an aircraft operated by the

[[Page 15]]

armed forces or an intelligence agency of the United States Government, 
or whose presence is required to perform, or is associated with the 
performance of, a governmental function.
    (2) An aircraft owned or operated by the armed forces or chartered 
to provide transportation to the armed forces if--
    (i) The aircraft is operated in accordance with title 10 of the 
United States Code;
    (ii) The aircraft is operated in the performance of a governmental 
function under title 14, 31, 32, or 50 of the United States Code and the 
aircraft is not used for commercial purposes; or
    (iii) The aircraft is chartered to provide transportation to the 
armed forces and the Secretary of Defense (or the Secretary of the 
department in which the Coast Guard is operating) designates the 
operation of the aircraft as being required in the national interest.
    (3) An aircraft owned or operated by the National Guard of a State, 
the District of Columbia, or any territory or possession of the United 
States, and that meets the criteria of paragraph (2) of this definition, 
qualifies as a public aircraft only to the extent that it is operated 
under the direct control of the Department of Defense.
    Rated 30-second OEI Power, with respect to rotorcraft turbine 
engines, means the approved brake horsepower developed under static 
conditions at specified altitudes and temperatures within the operating 
limitations established for the engine under part 33 of this chapter, 
for continuation of one flight operation after the failure or shutdown 
of one engine in multiengine rotorcraft, for up to three periods of use 
no longer than 30 seconds each in any one flight, and followed by 
mandatory inspection and prescribed maintenance action.
    Rated 2-minute OEI Power, with respect to rotorcraft turbine 
engines, means the approved brake horsepower developed under static 
conditions at specified altitudes and temperatures within the operating 
limitations established for the engine under part 33 of this chapter, 
for continuation of one flight operation after the failure or shutdown 
of one engine in multiengine rotorcraft, for up to three periods of use 
no longer than 2 minutes each in any one flight, and followed by 
mandatory inspection and prescribed maintenance action.
    Rated continuous OEI power, with respect to rotorcraft turbine 
engines, means the approved brake horsepower developed under static 
conditions at specified altitudes and temperatures within the operating 
limitations established for the engine under part 33 of this chapter, 
and limited in use to the time required to complete the flight after the 
failure or shutdown of one engine of a multiengine rotorcraft.
    Rated maximum continuous augmented thrust, with respect to turbojet 
engine type certification, means the approved jet thrust that is 
developed statically or in flight, in standard atmosphere at a specified 
altitude, with fluid injection or with the burning of fuel in a separate 
combustion chamber, within the engine operating limitations established 
under Part 33 of this chapter, and approved for unrestricted periods of 
use.
    Rated maximum continuous power, with respect to reciprocating, 
turbopropeller, and turboshaft engines, means the approved brake 
horsepower that is developed statically or in flight, in standard 
atmosphere at a specified altitude, within the engine operating 
limitations established under part 33, and approved for unrestricted 
periods of use.
    Rated maximum continuous thrust, with respect to turbojet engine 
type certification, means the approved jet thrust that is developed 
statically or in flight, in standard atmosphere at a specified altitude, 
without fluid injection and without the burning of fuel in a separate 
combustion chamber, within the engine operating limitations established 
under part 33 of this chapter, and approved for unrestricted periods of 
use.
    Rated takeoff augmented thrust, with respect to turbojet engine type 
certification, means the approved jet thrust that is developed 
statically under standard sea level conditions, with fluid injection or 
with the burning of fuel in a separate combustion chamber, within the 
engine operating limitations established under part 33 of this chapter, 
and limited in use to periods

[[Page 16]]

of not over 5 minutes for takeoff operation.
    Rated takeoff power, with respect to reciprocating, turbopropeller, 
and turboshaft engine type certification, means the approved brake 
horsepower that is developed statically under standard sea level 
conditions, within the engine operating limitations established under 
part 33, and limited in use to periods of not over 5 minutes for takeoff 
operation.
    Rated takeoff thrust, with respect to turbojet engine type 
certification, means the approved jet thrust that is developed 
statically under standard sea level conditions, without fluid injection 
and without the burning of fuel in a separate combustion chamber, within 
the engine operating limitations established under part 33 of this 
chapter, and limited in use to periods of not over 5 minutes for takeoff 
operation.
    Rated 30-minute OEI power, with respect to rotorcraft turbine 
engines, means the approved brake horsepower developed under static 
conditions at specified altitudes and temperatures within the operating 
limitations established for the engine under part 33 of this chapter, 
and limited in use to one period of use no longer than 30 minutes after 
the failure or shutdown of one engine of a multiengine rotorcraft.
    Rated 2\1/2\-minute OEI power, with respect to rotorcraft turbine 
engines, means the approved brake horsepower developed under static 
conditions at specified altitudes and temperatures within the operating 
limitations established for the engine under part 33 of this chapter for 
periods of use no longer than 2\1/2\ minutes each after the failure or 
shutdown of one engine of a multiengine rotorcraft.
    Rating means a statement that, as a part of a certificate, sets 
forth special conditions, privileges, or limitations.
    Reference landing speed means the speed of the airplane, in a 
specified landing configuration, at the point where it descends through 
the 50 foot height in the determination of the landing distance.
    Reporting point means a geographical location in relation to which 
the position of an aircraft is reported.
    Restricted area. A restricted area is airspace designated under Part 
73 within which the flight of aircraft, while not wholly prohibited, is 
subject to restriction.
    Rocket means an aircraft propelled by ejected expanding gases 
generated in the engine from self-contained propellants and not 
dependent on the intake of outside substances. It includes any part 
which becomes separated during the operation.
    Rotorcraft means a heavier-than-air aircraft that depends 
principally for its support in flight on the lift generated by one or 
more rotors.
    Rotorcraft-load combination means the combination of a rotorcraft 
and an external-load, including the external-load attaching means. 
Rotorcraft-load combinations are designated as Class A, Class B, Class 
C, and Class D, as follows:
    (1) Class A rotorcraft-load combination means one in which the 
external load cannot move freely, cannot be jettisoned, and does not 
extend below the landing gear.
    (2) Class B rotorcraft-load combination means one in which the 
external load is jettisonable and is lifted free of land or water during 
the rotorcraft operation.
    (3) Class C rotorcraft-load combination means one in which the 
external load is jettisonable and remains in contact with land or water 
during the rotorcraft operation.
    (4) Class D rotorcraft-load combination means one in which the 
external-load is other than a Class A, B, or C and has been specifically 
approved by the Administrator for that operation.
    Route segment is a portion of a route bounded on each end by a fix 
or navigation aid (NAVAID).
    Sea level engine means a reciprocating aircraft engine having a 
rated takeoff power that is producible only at sea level.
    Second in command means a pilot who is designated to be second in 
command of an aircraft during flight time.
    Show, unless the context otherwise requires, means to show to the 
satisfaction of the Administrator.
    Small aircraft means aircraft of 12,500 pounds or less, maximum 
certificated takeoff weight.
    Special VFR conditions mean meteorological conditions that are less 
than those required for basic VFR flight in

[[Page 17]]

controlled airspace and in which some aircraft are permitted flight 
under visual flight rules.
    Special VFR operations means aircraft operating in accordance with 
clearances within controlled airspace in meteorological conditions less 
than the basic VFR weather minima. Such operations must be requested by 
the pilot and approved by ATC.
    Standard atmosphere means the atmosphere defined in U.S. Standard 
Atmosphere, 1962 (Geopotential altitude tables).
    Stopway means an area beyond the takeoff runway, no less wide than 
the runway and centered upon the extended centerline of the runway, able 
to support the airplane during an aborted takeoff, without causing 
structural damage to the airplane, and designated by the airport 
authorities for use in decelerating the airplane during an aborted 
takeoff.
    Suitable RNAV system is an RNAV system that meets the required 
performance established for a type of operation, e.g. IFR; and is 
suitable for operation over the route to be flown in terms of any 
performance criteria (including accuracy) established by the air 
navigation service provider for certain routes (e.g. oceanic, ATS 
routes, and IAPs). An RNAV system's suitability is dependent upon the 
availability of ground and/or satellite navigation aids that are needed 
to meet any route performance criteria that may be prescribed in route 
specifications to navigate the aircraft along the route to be flown. 
Information on suitable RNAV systems is published in FAA guidance 
material.
    Synthetic vision means a computer-generated image of the external 
scene topography from the perspective of the flight deck that is derived 
from aircraft attitude, high-precision navigation solution, and database 
of terrain, obstacles and relevant cultural features.
    Synthetic vision system means an electronic means to display a 
synthetic vision image of the external scene topography to the flight 
crew.
    Takeoff power:
    (1) With respect to reciprocating engines, means the brake 
horsepower that is developed under standard sea level conditions, and 
under the maximum conditions of crankshaft rotational speed and engine 
manifold pressure approved for the normal takeoff, and limited in 
continuous use to the period of time shown in the approved engine 
specification; and
    (2) With respect to turbine engines, means the brake horsepower that 
is developed under static conditions at a specified altitude and 
atmospheric temperature, and under the maximum conditions of rotor shaft 
rotational speed and gas temperature approved for the normal takeoff, 
and limited in continuous use to the period of time shown in the 
approved engine specification.
    Takeoff safety speed means a referenced airspeed obtained after 
lift-off at which the required one-engine-inoperative climb performance 
can be achieved.
    Takeoff thrust, with respect to turbine engines, means the jet 
thrust that is developed under static conditions at a specific altitude 
and atmospheric temperature under the maximum conditions of rotorshaft 
rotational speed and gas temperature approved for the normal takeoff, 
and limited in continuous use to the period of time shown in the 
approved engine specification.
    Tandem wing configuration means a configuration having two wings of 
similar span, mounted in tandem.
    TCAS I means a TCAS that utilizes interrogations of, and replies 
from, airborne radar beacon transponders and provides traffic advisories 
to the pilot.
    TCAS II means a TCAS that utilizes interrogations of, and replies 
from airborne radar beacon transponders and provides traffic advisories 
and resolution advisories in the vertical plane.
    TCAS III means a TCAS that utilizes interrogation of, and replies 
from, airborne radar beacon transponders and provides traffic advisories 
and resolution advisories in the vertical and horizontal planes to the 
pilot.
    Time in service, with respect to maintenance time records, means the 
time from the moment an aircraft leaves the surface of the earth until 
it touches it at the next point of landing.
    Traffic pattern means the traffic flow that is prescribed for 
aircraft landing at, taxiing on, or taking off from, an airport.

[[Page 18]]

    True airspeed means the airspeed of an aircraft relative to 
undisturbed air. True airspeed is equal to equivalent airspeed 
multiplied by ([rho]0/[rho])\1/2\.
    Type:
    (1) As used with respect to the certification, ratings, privileges, 
and limitations of airmen, means a specific make and basic model of 
aircraft, including modifications thereto that do not change its 
handling or flight characteristics. Examples include: DC-7, 1049, and F-
27; and
    (2) As used with respect to the certification of aircraft, means 
those aircraft which are similar in design. Examples include: DC-7 and 
DC-7C; 1049G and 1049H; and F-27 and F-27F.
    (3) As used with respect to the certification of aircraft engines 
means those engines which are similar in design. For example, JT8D and 
JT8D-7 are engines of the same type, and JT9D-3A and JT9D-7 are engines 
of the same type.
    United States, in a geographical sense, means (1) the States, the 
District of Columbia, Puerto Rico, and the possessions, including the 
territorial waters, and (2) the airspace of those areas.
    United States air carrier means a citizen of the United States who 
undertakes directly by lease, or other arrangement, to engage in air 
transportation.
    VFR over-the-top, with respect to the operation of aircraft, means 
the operation of an aircraft over-the-top under VFR when it is not being 
operated on an IFR flight plan.
    Warning area. A warning area is airspace of defined dimensions, 
extending from 3 nautical miles outward from the coast of the United 
States, that contains activity that may be hazardous to nonparticipating 
aircraft. The purpose of such warning areas is to warn nonparticipating 
pilots of the potential danger. A warning area may be located over 
domestic or international waters or both.
    Weight-shift-control aircraft means a powered aircraft with a framed 
pivoting wing and a fuselage controllable only in pitch and roll by the 
pilot's ability to change the aircraft's center of gravity with respect 
to the wing. Flight control of the aircraft depends on the wing's 
ability to flexibly deform rather than the use of control surfaces.
    Winglet or tip fin means an out-of-plane surface extending from a 
lifting surface. The surface may or may not have control surfaces.

[Doc. No. 1150, 27 FR 4588, May 15, 1962]

    Editorial Note: For Federal Register citations affecting Sec. 1.1, 
see the List of CFR Sections Affected, which appears in the Finding Aids 
section of the printed volume and at www.fdsys.gov.



Sec. 1.2  Abbreviations and symbols.

    In Subchapters A through K of this chapter:
    AFM means airplane flight manual.
    AGL means above ground level.
    ALS means approach light system.
    APU means auxiliary power unit.
    ASR means airport surveillance radar.
    ATC means air traffic control.
    ATS means Air Traffic Service.
    CAMP means continuous airworthiness maintenance program.
    CAS means calibrated airspeed.
    CAT II means Category II.
    CHDO means an FAA Flight Standards certificate holding district 
office.
    CMP means configuration, maintenance, and procedures.
    CONSOL or CONSOLAN means a kind of low or medium frequency long 
range navigational aid.
    DH means decision height.
    DME means distance measuring equipment compatible with TACAN.
    EAS means equivalent airspeed.
    EFVS means enhanced flight vision system.
    Equi-Time Point means a point on the route of flight where the 
flight time, considering wind, to each of two selected airports is 
equal.
    ETOPS means extended operations.
    EWIS, as defined by Sec. 25.1701 of this chapter, means electrical 
wiring interconnection system.
    FAA means Federal Aviation Administration.
    FFS means full flight simulator.
    FM means fan marker.
    FSTD means flight simulation training device.
    FTD means flight training device.
    GS means glide slope.
    HIRL means high-intensity runway light system.

[[Page 19]]

    IAS means indicated airspeed.
    ICAO means International Civil Aviation Organization.
    IFR means instrument flight rules.
    IFSD means in-flight shutdown.
    ILS means instrument landing system.
    IM means ILS inner marker.
    INT means intersection.
    LDA means localizer-type directional aid.
    LFR means low-frequency radio range.
    LMM means compass locator at middle marker.
    LOC means ILS localizer.
    LOM means compass locator at outer marker.
    M means mach number.
    MAA means maximum authorized IFR altitude.
    MALS means medium intensity approach light system.
    MALSR means medium intensity approach light system with runway 
alignment indicator lights.
    MCA means minimum crossing altitude.
    MDA means minimum descent altitude.
    MEA means minimum en route IFR altitude.
    MEL means minimum equipment list.
    MM means ILS middle marker.
    MOCA means minimum obstruction clearance altitude.
    MRA means minimum reception altitude.
    MSL means mean sea level.
    NDB (ADF) means nondirectional beacon (automatic direction finder).
    NM means nautical mile.
    NOPAC means North Pacific area of operation.
    NOPT means no procedure turn required.
    OEI means one engine inoperative.
    OM means ILS outer marker.
    OPSPECS means operations specifications.
    PACOTS means Pacific Organized Track System.
    PAR means precision approach radar.
    PMA means parts manufacturer approval.
    PTRS means Performance Tracking and Reporting System.
    RAIL means runway alignment indicator light system.
    RBN means radio beacon.
    RCLM means runway centerline marking.
    RCLS means runway centerline light system.
    REIL means runway end identification lights.
    RFFS means rescue and firefighting services.
    RNAV means area navigation.
    RR means low or medium frequency radio range station.
    RVR means runway visual range as measured in the touchdown zone 
area.
    SALS means short approach light system.
    SATCOM means satellite communications.
    SSALS means simplified short approach light system.
    SSALSR means simplified short approach light system with runway 
alignment indicator lights.
    TACAN means ultra-high frequency tactical air navigational aid.
    TAS means true airspeed.
    TCAS means a traffic alert and collision avoidance system.
    TDZL means touchdown zone lights.
    TSO means technical standard order.
    TVOR means very high frequency terminal omnirange station.
    VA means design maneuvering speed.
    VB means design speed for maximum gust intensity.
    VC means design cruising speed.
    VD means design diving speed.
    VDF/MDF means demonstrated flight diving speed.
    VEF means the speed at which the critical engine is assumed to fail 
during takeoff.
    VF means design flap speed.
    VFC/MFC means maximum speed for stability characteristics.
    VFE means maximum flap extended speed.
    VFTO means final takeoff speed.
    VH means maximum speed in level flight with maximum continuous 
power.
    VLE means maximum landing gear extended speed.
    VLO means maximum landing gear operating speed.
    VLOF means lift-off speed.
    VMC means minimum control speed with the critical engine 
inoperative.

[[Page 20]]

    VMO/MMO means maximum operating limit speed.
    VMU means minimum unstick speed.
    VNE means never-exceed speed.
    VNO means maximum structural cruising speed.
    VR means rotation speed.
    VREF means reference landing speed.
    VS means the stalling speed or the minimum steady flight speed at 
which the airplane is controllable.
    VS0 means the stalling speed or the minimum steady flight 
speed in the landing configuration.
    VS1 means the stalling speed or the minimum steady flight 
speed obtained in a specific configuration.
    VSR means reference stall speed.
    VSRO means reference stall speed in the landing 
configuration.
    VSR1 means reference stall speed in a specific 
configuration.
    VSW means speed at which onset of natural or artificial stall 
warning occurs.
    VTOSS means takeoff safety speed for Category A rotorcraft.
    VX means speed for best angle of climb.
    VY means speed for best rate of climb.
    V1 means the maximum speed in the takeoff at which the pilot must 
take the first action (e.g., apply brakes, reduce thrust, deploy speed 
brakes) to stop the airplane within the accelerate-stop distance. 
V1 also means the minimum speed in the takeoff, following a 
failure of the critical engine at VEF, at which the pilot can 
continue the takeoff and achieve the required height above the takeoff 
surface within the takeoff distance.
    V2 means takeoff safety speed.
    V2min means minimum takeoff safety speed.
    VFR means visual flight rules.
    VHF means very high frequency.
    VOR means very high frequency omnirange station.
    VORTAC means collocated VOR and TACAN.

[Doc. No. 1150, 27 FR 4590, May 15, 1962]

    Editorial Note: For Federal Register citations affecting Sec. 1.2, 
see the List of CFR Sections Affected, which appears in the Finding Aids 
section of the printed volume and at www.fdsys.gov.



Sec. 1.3  Rules of construction.

    (a) In Subchapters A through K of this chapter, unless the context 
requires otherwise:
    (1) Words importing the singular include the plural;
    (2) Words importing the plural include the singular; and
    (3) Words importing the masculine gender include the feminine.
    (b) In Subchapters A through K of this chapter, the word:
    (1) Shall is used in an imperative sense;
    (2) May is used in a permissive sense to state authority or 
permission to do the act prescribed, and the words ``no person may * * 
*'' or ``a person may not * * *'' mean that no person is required, 
authorized, or permitted to do the act prescribed; and
    (3) Includes means ``includes but is not limited to''.

[Doc. No. 1150, 27 FR 4590, May 15, 1962, as amended by Amdt. 1-10, 31 
FR 5055, Mar. 29, 1966]



PART 3_GENERAL REQUIREMENTS--Table of Contents



Sec.
3.1 Applicability.
3.5 Statements about products, parts, appliances and materials.

    Authority: 49 U.S.C. 106(g), 40113, 44701, and 44704.

    Source: 70 FR 54832, Sept. 16, 2005, unless otherwise noted.



Sec. 3.1  Applicability.

    (a) This part applies to any person who makes a record regarding:
    (1) A type-certificated product, or
    (2) A product, part, appliance or material that may be used on a 
type-certificated product.
    (b) Section 3.5(b) does not apply to records made under part 43 of 
this chapter.



Sec. 3.5  Statements about products, parts, appliances and materials.

    (a) Definitions. The following terms will have the stated meanings 
when used in this section:
    Airworthy means the aircraft conforms to its type design and is in a 
condition for safe operation.

[[Page 21]]

    Product means an aircraft, aircraft engine, or aircraft propeller.
    Record means any writing, drawing, map, recording, tape, film, 
photograph or other documentary material by which information is 
preserved or conveyed in any format, including, but not limited to, 
paper, microfilm, identification plates, stamped marks, bar codes or 
electronic format, and can either be separate from, attached to or 
inscribed on any product, part, appliance or material.
    (b) Prohibition against fraudulent and intentionally false 
statements. When conveying information related to an advertisement or 
sales transaction, no person may make or cause to be made:
    (1) Any fraudulent or intentionally false statement in any record 
about the airworthiness of a type-certificated product, or the 
acceptability of any product, part, appliance, or material for 
installation on a type-certificated product.
    (2) Any fraudulent or intentionally false reproduction or alteration 
of any record about the airworthiness of any type-certificated product, 
or the acceptability of any product, part, appliance, or material for 
installation on a type-certificated product.
    (c) Prohibition against intentionally misleading statements. (1) 
When conveying information related to an advertisement or sales 
transaction, no person may make, or cause to be made, a material 
representation that a type-certificated product is airworthy, or that a 
product, part, appliance, or material is acceptable for installation on 
a type-certificated product in any record if that representation is 
likely to mislead a consumer acting reasonably under the circumstances.
    (2) When conveying information related to an advertisement or sales 
transaction, no person may make, or cause to be made, through the 
omission of material information, a representation that a type-
certificated product is airworthy, or that a product, part, appliance, 
or material is acceptable for installation on a type-certificated 
product in any record if that representation is likely to mislead a 
consumer acting reasonably under the circumstances.
    (d) The provisions of Sec. 3.5(b) and Sec. 3.5(c) shall not apply 
if a person can show that the product is airworthy or that the product, 
part, appliance or material is acceptable for installation on a type-
certificated product.

[[Page 22]]



                      SUBCHAPTER B_PROCEDURAL RULES





PART 11_GENERAL RULEMAKING PROCEDURES--Table of Contents



                     Subpart A_Rulemaking Procedures

Sec.
11.1 To what does this part apply?

                           Definition of Terms

11.3 What is an advance notice of proposed rulemaking?
11.5 What is a notice of proposed rulemaking?
11.7 What is a supplemental notice of proposed rulemaking?
11.9 What is a final rule?
11.11 What is a final rule with request for comments?
11.13 What is a direct final rule?
11.15 What is a petition for exemption?
11.17 What is a petition for rulemaking?
11.19 What is a special condition?

                                 General

11.21 What are the most common kinds of rulemaking actions for which FAA 
          follows the Administrative Procedure Act?
11.23 Does FAA follow the same procedures in issuing all types of rules?
11.25 How does FAA issue rules?
11.27 Are there other ways FAA collects specific rulemaking 
          recommendations before we issue an NPRM?
11.29 May FAA change its regulations without first issuing an ANPRM or 
          NPRM?
11.31 How does FAA process direct final rules?
11.33 How can I track FAA's rulemaking activities?
11.35 Does FAA include sensitive security information and proprietary 
          information in the Federal Docket Management System (FDMS)?
11.37 Where can I find information about an Airworthiness Directive, an 
          airspace designation, or a petition handled in a region?
11.38 What public comment procedures does FAA follow for Special 
          Conditions?
11.39 How may I participate in FAA's rulemaking process?
11.40 Can I get more information about a rulemaking?

                            Written Comments

11.41 Who may file comments?
11.43 What information must I put in my written comments?
11.45 Where and when do I file my comments?
11.47 May I ask for more time to file my comments?

                  Public Meetings and Other Proceedings

11.51 May I request that FAA hold a public meeting on a rulemaking 
          action?
11.53 What takes place at a public meeting?

               Petitions for Rulemaking and for Exemptions

11.61 May I ask FAA to adopt, amend, or repeal a regulation, or grant 
          relief from the requirements of a current regulation?
11.63 How and to whom do I submit my petition for rulemaking or petition 
          for exemption?
11.71 What information must I include in my petition for rulemaking?
11.73 How does FAA process petitions for rulemaking?
11.75 Does FAA invite public comment on petitions for rulemaking?
11.77 Is there any additional information I must include in my petition 
          for designating airspace?
11.81 What information must I include in my petition for an exemption?
11.83 How can I operate under an exemption outside the United States?
11.85 Does FAA invite public comment on petitions for exemption?
11.87 Are there circumstances in which FAA may decide not to publish a 
          summary of my petition for exemption?
11.89 How much time do I have to submit comments to FAA on a petition 
          for exemption?
11.91 How does FAA inform me of its decision on my petition for 
          exemption?
11.101 May I ask FAA to reconsider my petition for rulemaking or 
          petition for exemption if it is denied?
11.103 What exemption relief may be available to federal, state, and 
          local governments when operating aircraft that are not public 
          aircraft?

            Subpart B_Paperwork Reduction Act Control Numbers

11.201 Office of Management and Budget (OMB) control numbers assigned 
          under the Paperwork Reduction Act.

Appendix 1 to Part 11--Oral Communications With the Public During 
          Rulemaking

    Authority: 49 U.S.C. 106(f), 106(g), 40101, 40103, 40105, 40109, 
40113, 44110, 44502, 44701-44702, 44711, and 46102.

    Source: Docket No. 1999-6622, 65 FR 50863, Aug. 21, 2000, unless 
otherwise noted.

    Editorial Note: Nomenclature changes to part 11 appear at 61 FR 
18052, April 24, 1996.

[[Page 23]]



                     Subpart A_Rulemaking Procedures



Sec. 11.1  To what does this part apply?

    This part applies to the issuance, amendment, and repeal of any 
regulation for which FAA (``we'') follows public rulemaking procedures 
under the Administrative Procedure Act (``APA'') (5 U.S.C. 553).

                           Definition of Terms



Sec. 11.3  What is an advance notice of proposed rulemaking?

    An advance notice of proposed rulemaking (ANPRM) tells the public 
that FAA is considering an area for rulemaking and requests written 
comments on the appropriate scope of the rulemaking or on specific 
topics. An advance notice of proposed rulemaking may or may not include 
the text of potential changes to a regulation.



Sec. 11.5  What is a notice of proposed rulemaking?

    A notice of proposed rulemaking (NPRM) proposes FAA's specific 
regulatory changes for public comment and contains supporting 
information. It includes proposed regulatory text.



Sec. 11.7  What is a supplemental notice of proposed rulemaking?

    On occasion, FAA may decide that it needs more information on an 
issue, or that we should take a different approach than we proposed. 
Also, we may want to follow a commenter's suggestion that goes beyond 
the scope of the original proposed rule. In these cases, FAA may issue a 
supplemental notice of proposed rulemaking (SNPRM) to give the public an 
opportunity to comment further or to give us more information.



Sec. 11.9  What is a final rule?

    A final rule sets out new or revised requirements and their 
effective date. It also may remove requirements. When preceded by an 
NPRM, a final rule will also identify significant substantive issues 
raised by commenters in response to the NPRM and will give the agency's 
response.



Sec. 11.11  What is a final rule with request for comments?

    A final rule with request for comment is a rule that the FAA issues 
in final (with an effective date) that invites public comment on the 
rule. We usually do this when we have not first issued an ANPRM or NPRM, 
because we have found that doing so would be impracticable, unnecessary, 
or contrary to the public interest. We give our reasons for our 
determination in the preamble. The comment period often ends after the 
effective date of the rule. A final rule not preceded by an ANPRM or 
NPRM is commonly called an ``immediately adopted final rule.'' We invite 
comments on these rules only if we think that we will receive useful 
information. For example, we would not invite comments when we are just 
making an editorial clarification or correction.



Sec. 11.13  What is a direct final rule?

    A direct final rule is a type of final rule with request for 
comments. Our reason for issuing a direct final rule without an NPRM is 
that we would not expect to receive any adverse comments, and so an NPRM 
is unnecessary. However, to be certain that we are correct, we set the 
comment period to end before the effective date. If we receive an 
adverse comment or notice of intent to file an adverse comment, we then 
withdraw the final rule before it becomes effective and may issue an 
NPRM.



Sec. 11.15  What is a petition for exemption?

    A petition for exemption is a request to FAA by an individual or 
entity asking for relief from the requirements of a current regulation.



Sec. 11.17  What is a petition for rulemaking?

    A petition for rulemaking is a request to FAA by an individual or 
entity asking the FAA to adopt, amend, or repeal a regulation.



Sec. 11.19  What is a special condition?

    A special condition is a regulation that applies to a particular 
aircraft design. The FAA issues special conditions when we find that the 
airworthiness

[[Page 24]]

regulations for an aircraft, aircraft engine, or propeller design do not 
contain adequate or appropriate safety standards, because of a novel or 
unusual design feature.

                                 General



Sec. 11.21  What are the most common kinds of rulemaking actions
for which FAA follows the Administrative Procedure Act?

    FAA follows the Administrative Procedure Act (APA) procedures for 
these common types of rules:
    (a) Rules found in the Code of Federal Regulations;
    (b) Airworthiness directives issued under part 39 of this chapter; 
and
    (c) Airspace Designations issued under various parts of this 
chapter.



Sec. 11.23  Does FAA follow the same procedures in issuing all types
of rules?

    Yes, in general, FAA follows the same procedures for all rule types. 
There are some differences as to which FAA official has authority to 
issue each type, and where you send petitions for FAA to adopt, amend, 
or repeal each type. Assume that the procedures in this subpart apply to 
all rules, except where we specify otherwise.



Sec. 11.25  How does FAA issue rules?

    (a) The FAA uses APA rulemaking procedures to adopt, amend, or 
repeal regulations. To propose or adopt a new regulation, or to change a 
current regulation, FAA will issue one or more of the following 
documents. We publish these rulemaking documents in the Federal Register 
unless we name and personally serve a copy of a rule on every person 
subject to it. We also make all documents available to the public by 
posting them in the Federal Docket Management System at http://
www.regulations.gov.
    (1) An advance notice of proposed rulemaking (ANPRM).
    (2) A notice of proposed rulemaking (NPRM).
    (3) A supplemental notice of proposed rulemaking (SNPRM).
    (4) A final rule.
    (5) A final rule with request for comments.
    (6) A direct final rule.
    (b) Each of the rulemaking documents in paragraph (a) of this 
section generally contains the following information:
    (1) The topic involved in the rulemaking document.
    (2) FAA's legal authority for issuing the rulemaking document.
    (3) How interested persons may participate in the rulemaking 
proceeding (for example, by filing written comments or making oral 
presentations at a public meeting).
    (4) Whom to call if you have questions about the rulemaking 
document.
    (5) The date, time, and place of any public meetings FAA will hold 
to discuss the rulemaking document.
    (6) The docket number and regulation identifier number (RIN) for the 
rulemaking proceeding.

[Doc. No. 1999-6622, 65 FR 50863, Aug. 21, 2000, as amended at 72 FR 
68474, Dec. 5, 2007]



Sec. 11.27  Are there other ways FAA collects specific rulemaking
recommendations before we issue an NPRM?

    Yes, the FAA obtains advice and recommendations from rulemaking 
advisory committees. One of these committees is the Aviation Rulemaking 
Advisory Committee (ARAC), which is a formal standing committee 
comprised of representatives of aviation associations and industry, 
consumer groups, and interested individuals. In conducting its 
activities, ARAC complies with the Federal Advisory Committee Act and 
the direction of FAA. We task ARAC with providing us with recommended 
rulemaking actions dealing with specific areas and problems. If we 
accept an ARAC recommendation to change an FAA rule, we ordinarily 
publish an NPRM using the procedures in this part. The FAA may establish 
other rulemaking advisory committees as needed to focus on specific 
issues for a limited period of time.



Sec. 11.29  May FAA change its regulations without first issuing
an ANPRM or NPRM?

    The FAA normally adds or changes a regulation by issuing a final 
rule after an NPRM. However, FAA may adopt, amend, or repeal regulations 
without

[[Page 25]]

first issuing an ANPRM or NPRM in the following situations:
    (a) We may issue a final rule without first requesting public 
comment if, for good cause, we find that an NPRM is impracticable, 
unnecessary, or contrary to the public interest. We place that finding 
and a brief statement of the reasons for it in the final rule. For 
example, we may issue a final rule in response to a safety emergency.
    (b) If an NPRM would be unnecessary because we do not expect to 
receive adverse comment, we may issue a direct final rule.



Sec. 11.31  How does FAA process direct final rules?

    (a) A direct final rule will take effect on a specified date unless 
FAA receives an adverse comment or notice of intent to file an adverse 
comment within the comment period--generally 60 days after the direct 
final rule is published in the Federal Register. An adverse comment 
explains why a rule would be inappropriate, or would be ineffective or 
unacceptable without a change. It may challenge the rule's underlying 
premise or approach. Under the direct final rule process, we do not 
consider the following types of comments to be adverse:
    (1) A comment recommending another rule change, in addition to the 
change in the direct final rule at issue. We consider the comment 
adverse, however, if the commenter states why the direct final rule 
would be ineffective without the change.
    (2) A frivolous or insubstantial comment.
    (b) If FAA has not received an adverse comment or notice of intent 
to file an adverse comment, we will publish a confirmation document in 
the Federal Register, generally within 15 days after the comment period 
closes. The confirmation document tells the public the effective date of 
the rule.
    (c) If we receive an adverse comment or notice of intent to file an 
adverse comment, we will advise the public by publishing a document in 
the Federal Register before the effective date of the direct final rule. 
This document may withdraw the direct final rule in whole or in part. If 
we withdraw a direct final rule because of an adverse comment, we may 
incorporate the commenter's recommendation into another direct final 
rule or may publish a notice of proposed rulemaking.



Sec. 11.33  How can I track FAA's rulemaking activities?

    The best ways to track FAA's rulemaking activities are with the 
docket number or the regulation identifier number.
    (a) Docket ID. We assign a docket ID to each rulemaking document 
proceeding. Each rulemaking document FAA issues in a particular 
rulemaking proceeding, as well as public comments on the proceeding, 
will display the same docket ID. This ID allows you to search the 
Federal Docket Management System (FDMS) for information on most 
rulemaking proceedings. You can view and copy docket materials during 
regular business hours at the U.S. Department of Transportation, Docket 
Operations, West Building Ground Floor, Room W12-140, 1200 New Jersey 
Avenue, SE., Washington, DC 20590. Or you can view and download docketed 
materials through the Internet at http://www.regulations.gov. If you 
can't find the material in the electronic docket, contact the person 
listed under FOR FURTHER INFORMATION CONTACT in the document you are 
interested in.
    (b) Regulation identifier number. DOT publishes a semiannual agenda 
of all current and projected DOT rulemakings, reviews of existing 
regulations, and completed actions. This semiannual agenda appears in 
the Unified Agenda of Federal Regulations, published in the Federal 
Register in April and October of each year. The semiannual agenda tells 
the public about DOT's--including FAA's--regulatory activities. DOT 
assigns a regulation identifier number (RIN) to each individual 
rulemaking proceeding in the semiannual agenda. This number appears on 
all rulemaking documents published in the Federal Register and makes it 
easy for you to track those rulemaking proceedings in both the Federal 
Register and the semiannual regulatory agenda.

[Doc. No. 1999-6622, 65 FR 50863, Aug. 21, 2000, as amended at 72 FR 
68474, Dec. 5, 2007]

[[Page 26]]



Sec. 11.35  Does FAA include sensitive security information and
proprietary information in the Federal Docket Management System (FDMS)?

    (a) Sensitive security information. You should not submit sensitive 
security information to the rulemaking docket, unless you are invited to 
do so in our request for comments. If we ask for this information, we 
will tell you in the specific document how to submit this information, 
and we will provide a separate non-public docket for it. For all 
proposed rule changes involving civil aviation security, we review 
comments as we receive them, before they are placed in the docket. If we 
find that a comment contains sensitive security information, we remove 
that information before placing the comment in the general docket.
    (b) Proprietary information. When we are aware of proprietary 
information filed with a comment, we do not place it in the docket. We 
hold it in a separate file to which the public does not have access, and 
place a note in the docket that we have received it. If we receive a 
request to examine or copy this information, we treat it as any other 
request under the Freedom of Information Act (5 U.S.C. 552). We process 
such a request under the DOT procedures found in 49 CFR part 7.

[Doc. No. 1999-6622, 65 FR 50863, Aug. 21, 2000, as amended at 72 FR 
68474, Dec. 5, 2007]



Sec. 11.37  Where can I find information about an Airworthiness
Directive, an airspace designation, or a petition handled in a region?

    The FAA includes most documents concerning Airworthiness Directives, 
airspace designations, or petitions handled in a region in the 
electronic docket. If the information isn't in the docket, contact the 
person listed under FOR FURTHER INFORMATION CONTACT in the Federal 
Register document about the action.



Sec. 11.38  What public comment procedures does the FAA follow for
Special Conditions?

    Even though the Administrative Procedure Act does not require notice 
and comment for rules of particular applicability, FAA does publish 
proposed special conditions for comment. In the following circumstances 
we may not invite comment before we issue a special condition. If we 
don't, we will invite comment when we publish the final special 
condition.
    (a) The FAA considers prior notice to be impracticable if issuing a 
design approval would significantly delay delivery of the affected 
aircraft. We consider such a delay to be contrary to the public 
interest.
    (b) The FAA considers prior notice to be unnecessary if we have 
provided previous opportunities to comment on substantially identical 
proposed special conditions, and we are satisfied that new comments are 
unlikely.



Sec. 11.39  How may I participate in FAA's rulemaking process?

    You may participate in FAA's rulemaking process by doing any of the 
following:
    (a) File written comments on any rulemaking document that asks for 
comments, including an ANPRM, NPRM, SNPRM, a final rule with request for 
comments, or a direct final rule. Follow the directions for commenting 
found in each rulemaking document.
    (b) Ask that we hold a public meeting on any rulemaking, and 
participate in any public meeting that we hold.
    (c) File a petition for rulemaking that asks us to adopt, amend, or 
repeal a regulation.



Sec. 11.40  Can I get more information about a rulemaking?

    You can contact the person listed under FOR FURTHER INFORMATION 
CONTACT in the preamble of a rule. That person can explain the meaning 
and intent of a proposed rule, the technical aspects of a document, the 
terminology in a document, and can tell you our published schedule for 
the rulemaking process. We cannot give you information that is not 
already available to other members of the public. Department of 
Transportation policy on oral communications with the public during 
rulemaking appears in appendix 1 of this part.

[[Page 27]]

                            Written Comments



Sec. 11.41  Who may file comments?

    Anyone may file written comments about proposals and final rules 
that request public comments.



Sec. 11.43  What information must I put in my written comments?

    (a) Your written comments must be in English and must contain the 
following:
    (1) The docket number of the rulemaking document you are commenting 
on, clearly set out at the beginning of your comments.
    (2) Your name and mailing address, and, if you wish, other contact 
information, such as a fax number, telephone number, or e-mail address.
    (3) Your information, views, or arguments, following the 
instructions for participation in the rulemaking document on which you 
are commenting.
    (b) You should also include all material relevant to any statement 
of fact or argument in your comments, to the extent that the material is 
available to you and reasonable for you to submit. Include a copy of the 
title page of the document. Whether or not you submit a copy of the 
material to which you refer, you should indicate specific places in the 
material that support your position.



Sec. 11.45  Where and when do I file my comments?

    (a) Send your comments to the location specified in the rulemaking 
document on which you are commenting. If you are asked to send your 
comments to the Federal Document Management System, you may send them in 
either of the following ways:
    (1) By mail to: U.S. Department of Transportation, Docket 
Operations, West Building Ground Floor, Room W12-140, 1200 New Jersey 
Avenue, SE., Washington, DC 20590.
    (2) Through the Internet to http://www.regulations.gov.
    (3) In any other manner designated by FAA.
    (b) Make sure that your comments reach us by the deadline set out in 
the rulemaking document on which you are commenting. We will consider 
late-filed comments to the extent possible only if they do not 
significantly delay the rulemaking process.
    (c) We may reject your paper or electronic comments if they are 
frivolous, abusive, or repetitious. We may reject comments you file 
electronically if you do not follow the electronic filing instructions 
at the Federal Docket Management System Web site.

[Doc. No. 1999-6622, 65 FR 50863, Aug. 21, 2000, as amended at 72 FR 
68474, Dec. 5, 2007]



Sec. 11.47  May I ask for more time to file my comments?

    Yes, if FAA grants your request for more time to file comments, we 
grant all persons the same amount of time. We will notify the public of 
the extension by a document in the Federal Register. If FAA denies your 
request, we will notify you of the denial. To ask for more time, you 
must file a written or electronic request for extension at least 10 days 
before the end of the comment period. Your letter or message must--
    (a) Show the docket number of the rule at the top of the first page;
    (b) State, at the beginning, that you are requesting an extension of 
the comment period;
    (c) Show that you have good cause for the extension and that an 
extension is in the public interest;
    (d) Be sent to the address specified for comments in the rulemaking 
document on which you are commenting.

                  Public Meetings and Other Proceedings



Sec. 11.51  May I request that FAA hold a public meeting on a
rulemaking action?

    Yes, you may request that we hold a public meeting. FAA holds a 
public meeting when we need more than written comments to make a fully 
informed decision. Submit your written request to the address specified 
in the rulemaking document on which you are commenting. Specify at the 
top of your letter or message that you are requesting that the agency 
hold a public meeting. Submit your request no later than 30 days after 
our rulemaking notice. If we find good cause for a meeting, we will 
notify you and publish a

[[Page 28]]

notice of the meeting in the Federal Register.



Sec. 11.53  What takes place at a public meeting?

    A public meeting is a non-adversarial, fact-finding proceeding 
conducted by an FAA representative. Public meetings are announced in the 
Federal Register. We invite interested persons to attend and to present 
their views to the agency on specific issues. There are no formal 
pleadings and no adverse parties, and any regulation issued afterward is 
not necessarily based exclusively on the record of the meeting.

               Petitions for Rulemaking and for Exemption



Sec. 11.61  May I ask FAA to adopt, amend, or repeal a regulation,
or grant relief from the requirements of a current regulation?

    (a) Using a petition for rulemaking, you may ask FAA to add a new 
regulation to title 14 of the Code of Federal Regulations (14 CFR) or 
ask FAA to amend or repeal a current regulation in 14 CFR.
    (b) Using a petition for exemption, you may ask FAA to grant you 
relief from current regulations in 14 CFR.



Sec. 11.63  How and to whom do I submit my petition for rulemaking
or petition for exemption?

    (a) To submit a petition for rulemaking or exemption--
    (1) By electronic submission, submit your petition for rulemaking or 
exemption to FAA through the Internet at http://www.regulations.gov, the 
Federal Docket Management System Web site. For additional instructions, 
you may visit http://www.faa.gov/regulations.
    (2) By paper submission, send the original signed copy of your 
petition for rulemaking or exemption to this address: U.S. Department of 
Transportation, Docket Operations, West Building Ground Floor, Room W12-
140, 1200 New Jersey Avenue, SE., Washington, DC 20590.
    (b) Submit a petition for rulemaking or exemption from part 139 of 
this chapter--
    (1) To the appropriate FAA airport field office in whose area your 
airport is, or will be, established; and
    (2) To the U.S. Department of Transportation, Docket Operations, 
West Building Ground Floor, Room W12-140, 1200 New Jersey Avenue, SE., 
Washington, DC 20590 or by electronic submission to this Internet 
address: http://www.regulations.gov.
    (c) The FAA may designate other means by which you can submit 
petitions in the future.
    (d) Submit your petition for exemption 120 days before you need the 
exemption to take effect.

[Amdt. 11-50, 69 FR 22386, Apr. 26, 2004, as amended at 72 FR 68474, 
Dec. 5, 2007; Amdt. 11-55, 74 FR 202, Jan. 5, 2009]



Sec. 11.71  What information must I include in my petition for 
rulemaking?

    (a) You must include the following information in your petition for 
rulemaking:
    (1) Your name and mailing address and, if you wish, other contact 
information such as a fax number, telephone number, or e-mail address.
    (2) An explanation of your proposed action and its purpose.
    (3) The language you propose for a new or amended rule, or the 
language you would remove from a current rule.
    (4) An explanation of why your proposed action would be in the 
public interest.
    (5) Information and arguments that support your proposed action, 
including relevant technical and scientific data available to you.
    (6) Any specific facts or circumstances that support or demonstrate 
the need for the action you propose.
    (b) In the process of considering your petition, we may ask that you 
provide information or data available to you about the following:
    (1) The costs and benefits of your proposed action to society in 
general, and identifiable groups within society in particular.
    (2) The regulatory burden of your proposed action on small 
businesses, small organizations, small governmental jurisdictions, and 
Indian tribes.

[[Page 29]]

    (3) The recordkeeping and reporting burdens of your proposed action 
and whom the burdens would affect.
    (4) The effect of your proposed action on the quality of the natural 
and social environments.



Sec. 11.73  How does FAA process petitions for rulemaking?

    After we have determined the disposition of your petition, we will 
contact you in writing about our decision. The FAA may respond to your 
petition for rulemaking in one of the following ways:
    (a) If we determine that your petition justifies our taking the 
action you suggest, we may issue an NPRM or ANPRM. We will do so no 
later than 6 months after the date we receive your petition. In making 
our decision, we consider:
    (1) The immediacy of the safety or security concerns you raise;
    (2) The priority of other issues the FAA must deal with; and
    (3) The resources we have available to address these issues.
    (b) If we have issued an ANPRM or NPRM on the subject matter of your 
petition, we will consider your arguments for a rule change as a comment 
in connection with the rulemaking proceeding. We will not treat your 
petition as a separate action.
    (c) If we have begun a rulemaking project in the subject area of 
your petition, we will consider your comments and arguments for a rule 
change as part of that project. We will not treat your petition as a 
separate action.
    (d) If we have tasked ARAC to study the general subject area of your 
petition, we will ask ARAC to review and evaluate your proposed action. 
We will not treat your petition as a separate action.
    (e) If we determine that the issues you identify in your petition 
may have merit, but do not address an immediate safety concern or cannot 
be addressed because of other priorities and resource constraints, we 
may dismiss your petition. Your comments and arguments for a rule change 
will be placed in a database, which we will examine when we consider 
future rulemaking.



Sec. 11.75  Does FAA invite public comment on petitions for rulemaking?

    Generally, FAA does not invite public comment on petitions for 
rulemaking.



Sec. 11.77  Is there any additional information I must include in my
petition for designating airspace?

    In petitions asking FAA to establish, amend, or repeal a designation 
of airspace, including special use airspace, you must include all the 
information specified by Sec. 11.71 and also:
    (a) The location and a description of the airspace you want assigned 
or designated;
    (b) A complete description of the activity or use to be made of that 
airspace, including a detailed description of the type, volume, 
duration, time, and place of the operations to be conducted in the area;
    (c) A description of the air navigation, air traffic control, 
surveillance, and communication facilities available and to be provided 
if we grant the designation; and
    (d) The name and location of the agency, office, facility, or person 
who would have authority to permit the use of the airspace when it was 
not in use for the purpose to which you want it assigned.



Sec. 11.81  What information must I include in my petition for an 
exemption?

    You must include the following information in your petition for an 
exemption and submit it to FAA as soon as you know you need an 
exemption.
    (a) Your name and mailing address and, if you wish, other contact 
information such as a fax number, telephone number, or e-mail address;
    (b) The specific section or sections of 14 CFR from which you seek 
an exemption;
    (c) The extent of relief you seek, and the reason you seek the 
relief;
    (d) The reasons why granting your request would be in the public 
interest; that is, how it would benefit the public as a whole;
    (e) The reasons why granting the exemption would not adversely 
affect safety, or how the exemption would provide a level of safety at 
least equal

[[Page 30]]

to that provided by the rule from which you seek the exemption;
    (f) A summary we can publish in the Federal Register, stating:
    (1) The rule from which you seek the exemption; and
    (2) A brief description of the nature of the exemption you seek;
    (g) Any additional information, views or arguments available to 
support your request; and
    (h) If you want to exercise the privileges of your exemption outside 
the United States, the reason why you need to do so.



Sec. 11.83  How can I operate under an exemption outside the 
United States?

    If you want to be able to operate under your exemption outside the 
United States, you must request this when you petition for relief and 
give us the reason for this use. If you do not provide your reason or we 
determine that it does not justify this relief, we will limit your 
exemption to use within the United States. Before we extend your 
exemption for use outside the United States, we will verify that the 
exemption would be in compliance with the Standards of the International 
Civil Aviation Organization (ICAO). If it would not, but we still 
believe it would be in the public interest to allow you to do so, we 
will file a difference with ICAO. However, a foreign country still may 
not allow you to operate in that country without meeting the ICAO 
standard.



Sec. 11.85  Does FAA invite public comment on petitions for exemption?

    Yes, FAA publishes information about petitions for exemption in the 
Federal Register. The information includes--
    (a) The docket number of the petition;
    (b) The citation to the rule or rules from which the petitioner 
requested relief;
    (c) The name of the petitioner;
    (d) The petitioner's summary of the action requested and the reasons 
for requesting it; and
    (e) A request for comments to assist FAA in evaluating the petition.



Sec. 11.87  Are there circumstances in which FAA may decide not to
publish a summary of my petition for exemption?

    The FAA may not publish a summary of your petition for exemption and 
request comments if you present or we find good cause why we should not 
delay action on your petition. The factors we consider in deciding not 
to request comment include:
    (a) Whether granting your petition would set a precedent.
    (b) Whether the relief requested is identical to exemptions granted 
previously.
    (c) Whether our delaying action on your petition would affect you 
adversely.
    (d) Whether you filed your petition in a timely manner.



Sec. 11.89  How much time do I have to submit comments to FAA on a
petition for exemption?

    The FAA states the specific time allowed for comments in the Federal 
Register notice about the petition. We usually allow 20 days to comment 
on a petition for exemption.



Sec. 11.91  How does FAA inform me of its decision on my petition
for exemption?

    The FAA will notify you in writing about its decision on your 
petition. A copy of this decision is also placed in the public docket. 
We will include the docket number associated with your petition in our 
letter to you.

[Doc. No. FAA-2005-22982, 71 FR 1485, Jan. 10, 2006]



Sec. 11.101  May I ask FAA to reconsider my petition for rulemaking 
or petition for exemption if it is denied?

    Yes, you may petition FAA to reconsider your petition denial. You 
must submit your request to the address to which you sent your original 
petition, and FAA must receive it within 60 days after we issued the 
denial. For us to accept your petition, show the following:
    (a) That you have a significant additional fact and why you did not 
present it in your original petition;
    (b) That we made an important factual error in our denial of your 
original petition; or

[[Page 31]]

    (c) That we did not correctly interpret a law, regulation, or 
precedent.



Sec. 11.103  What exemption relief may be available to federal, state,
and local governments when operating aircraft that are not public
aircraft?

    The Federal Aviation Administration may grant a federal, state, or 
local government an exemption from part A of subtitle VII of title 49 
United States Code, and any regulation issued under that authority that 
is applicable to an aircraft as a result of the Independent Safety Board 
Act Amendments of 1994, Public Law 103-411, if--
    (a) The Administrator finds that granting the exemption is necessary 
to prevent an undue economic burden on the unit of government; and
    (b) The Administrator certifies that the aviation safety program of 
the unit of government is effective and appropriate to ensure safe 
operations of the type of aircraft operated by the unit of government.

[68 FR 25488, May 13, 2003]



            Subpart B_Paperwork Reduction Act Control Numbers



Sec. 11.201  Office of Management and Budget (OMB) control numbers
assigned under the Paperwork Reduction Act.

    (a) The Paperwork Reduction Act of 1995 (44 U.S.C. 3501-3520) 
requires FAA to get approval from OMB for our information collection 
activities, and to list a record of those approvals in the Federal 
Register. This subpart lists the control numbers OMB assigned to FAA's 
information collection activities.
    (b) The table listing OMB control numbers assigned to FAA's 
information collection activities follows:

------------------------------------------------------------------------
 14 CFR part or section identified and
               described                    Current OMB control number
------------------------------------------------------------------------
Part 14................................  2120-0539
Part 17................................  2120-0632
Part 21................................  2120-0018, 2120-0552
Part 34................................  2120-0508
Part 39................................  2120-0056
Part 43................................  2120-0020
Part 45................................  2120-0508
Part 47................................  2120-0024, 2120-0042
Part 49................................  2120-0043
Part 60................................  2120-0680
Part 61................................  2120-0021, 2120-0034, 2120-
                                          0543, 2120-0571
Part 63................................  2120-0007
Part 65................................  2120-0022, 2120-0535, 2120-
                                          0571, 2120-0648
Part 67................................  2120-0034, 2120-0543
Part 77................................  2120-0001
Part 91................................  2120-0005, 2120-0026, 2120-
                                          0027, 2120-0573, 2120-0606,
                                          2120-0620, 2120-0631, 2120-
                                          0651
Part 93................................  2120-0524, 2120-0606, 2120-0639
Part 101...............................  2120-0027
Part 105...............................  2120-0027, 2120-0641
Part 107...............................  2120-0075, 2120-0554, 2120-0628
Part 108...............................  2120-0098, 2120-0554, 2120-
                                          0577, 2120-0628, 2120-0642
Part 109...............................  2120-0505
Part 119...............................  2120-0593
Part 121...............................  2120-0008, 2120-0028, 2120-
                                          0535, 2120-0571, 2120-0600,
                                          2120-0606, 2120-0614, 2120-
                                          0616, 2120-0631, 2120-0651,
                                          2120-0653, 2120-0691, 2120-
                                          0702, 2120-0739
Part 125...............................  2120-0028, 2120-0085, 2120-
                                          0616, 2120-0651
Part 129...............................  2120-0028, 2120-0536, 2120-
                                          0616, 2120-0638
Part 133...............................  2120-0044
Part 135...............................  2120-0003, 2120-0028, 2120-
                                          0039, 2120-0535, 2120-0571,
                                          2120-0600, 2120-0606, 2120-
                                          0614, 2120-0616, 2120-0620,
                                          2120-0631, 2120-0653
Part 137...............................  2120-0049
Part 139...............................  2120-0045, 2120-0063
Part 141...............................  2120-0009
Part 142...............................  2120-0570
Part 145...............................  2120-0003, 2120-0010, 2120-0571
Part 147...............................  2120-0040
Part 150...............................  2120-0517
Part 157...............................  2120-0036
Part 158...............................  2120-0557
Part 161...............................  2120-0563
Part 171...............................  2120-0014
Part 183...............................  2120-0033, 2120-0604
Part 193...............................  2120-0646
Part 198...............................  2120-0514
Part 400...............................  2120-0643, 2120-0644, 0649
Part 401...............................  2120-0608
Part 440...............................  2120-0601
SFAR 36................................  2120-0507
SFAR 71................................  2120-0620
------------------------------------------------------------------------


[Doc. No. 1999-6622, 65 FR 50863, Aug. 21, 2000, as amended by Amdt. 11-
47, 67 FR 9553, Mar. 1, 2002; Amdt. 11-49, 68 FR 61321, Oct. 27, 2003; 
Amdt. 11-49, 68 FR 70132, Dec. 17, 2003; 70 FR 40163, July 12, 2005; 71 
FR 63426, Oct. 30, 2006; 72 FR 59599, Oct. 22, 2007; Amdt. 11-56, 79 FR 
12937, Mar. 7, 2014]

[[Page 32]]



 Sec. Appendix 1 to Part 11--Oral Communications With the Public During 
                               Rulemaking

                     1. What is an ex parte contact?

    ``Ex parte'' is a Latin term that means ``one sided,'' and indicates 
that not all parties to an issue were present when it was discussed. An 
ex parte contact involving rulemaking is any communication between FAA 
and someone outside the government regarding a specific rulemaking 
proceeding, before that proceeding closes. A rulemaking proceeding does 
not close until we publish the final rule or withdraw the NPRM. Because 
an ex parte contact excludes other interested persons, including the 
rest of the public, from the communication, it may give an unfair 
advantage to one party, or appear to do so.

        2. Are written comments to the docket ex parte contacts?

    Written comments submitted to the docket are not ex parte contacts 
because they are available for inspection by all members of the public.

               3. What is DOT policy on ex parte contacts?

    It is DOT policy to provide for open development of rules and to 
encourage full public participation in rulemaking actions. In addition 
to providing opportunity to respond in writing to an NPRM and to appear 
and be heard at a hearing, DOT policy encourages agencies to contact the 
public directly when we need factual information to resolve questions of 
substance. It also encourages DOT agencies to be receptive to 
appropriate contacts from persons affected by or interested in a 
proposed action. But under some circumstances an ex parte contact could 
affect the basic openness and fairness of the rulemaking process. Even 
the appearance of impropriety can affect public confidence in the 
process. For this reason, DOT policy sets careful guidelines for these 
contacts. The kind of ex parte contacts permitted and the procedures we 
follow depend on when the contact occurs in the rulemaking process.

  4. What kinds of ex parte contacts does DOT policy permit before we 
 issue an ANPRM, NPRM, Supplemental NPRM, or immediately adopted final 
                                  rule?

    The DOT policy authorizes ex parte contacts that we need to obtain 
technical and economic information. We need this information to decide 
whether to issue a regulation and what it should say. Each contact that 
influences our development of the regulation is noted in the preamble. 
For multiple contacts that are similar, we may provide only a general 
discussion. For contacts not discussed in the preamble, we place a 
report discussing each contact or group of related contacts in the 
rulemaking docket when it is opened.

 5. Does DOT policy permit ex parte contacts during the comment period?

    No, during the comment period, the public docket is available for 
written comments from any member of the public. These comments can be 
examined and responded to by any interested person. Because this public 
forum is available, DOT policy discourages ex parte contacts during the 
comment period. They are not necessary to collect the information the 
agency needs to make its decision.

6. What if the FAA believes it needs to meet with members of the public 
                        to discuss the proposal?

    If the FAA determines that it would be helpful to invite members of 
the public to make oral presentations to it regarding the proposal, we 
will announce a public meeting in the Federal Register.

 7. Are any oral contacts concerning the proposal permitted during the 
                             comment period?

    If you contact the agency with questions regarding the proposal 
during the comment period, we can only provide you with information that 
has already been made available to the general public. If you contact 
the agency to discuss the proposal, you will be told that the proper 
avenue of communication during the comment period is a written 
communication to the docket.

   8. If a substantive ex parte contact does occur during the comment 
                        period, what does FAA do?

    While FAA tries to ensure that FAA personnel and the public are 
aware of DOT policy, substantive ex parte contacts do occasionally 
occur, for example, at meetings not intended for that purpose. In such a 
case, we place a summary of the contact and a copy of any materials 
provided at the meeting in the rulemaking docket. We encourage 
participants in such a meeting to file written comments in the docket.

   9. Does DOT policy permit ex parte contacts the comment period has 
                                 closed?

    DOT policy strongly discourages ex parte contacts initiated by 
commenters to discuss their position on the proposal once the comment 
period has closed. Such a contact at this time would be improper, since 
other interested persons would not have an opportunity to respond. If we 
need further information regarding a comment in the docket, we may 
request this from a commenter. A record of this contact and the 
information provided is placed in the docket. If we need to make other 
contacts to update factual information, such as economic data, we will 
disclose this information in the final rule

[[Page 33]]

docket or in the economic studies accompanying it, which are available 
in the docket.

  10. What if FAA needs to meet with interested persons to discuss the 
              proposal after the comment period has closed?

    If FAA determines that it would be helpful to meet with a person or 
group after the close of the comment period to discuss a course of 
action to be taken, we will announce the meeting in the Federal 
Register. We will also consider reopening the comment period. If an 
inappropriate ex parte contact does occur after the comment period 
closes, a summary of the contact and a copy of any material distributed 
during meeting will be placed in the docket if it could be seen as 
influencing the rulemaking process.

    11. Under what circumstances will FAA reopen the comment period?

    If we receive an ex parte communication after the comment period has 
closed that could substantially influence the rulemaking, we may reopen 
the comment period. DOT policy requires the agency to carefully consider 
whether the substance of the contact will give the commenter an unfair 
advantage, since the rest of the public may not see the record of the 
contact in the docket. When the substance of a proposed rule is 
significantly changed as a result of such an oral communication, DOT 
policy and practice requires that the comment period be reopened by 
issuing a supplemental NPRM in which the reasons for the change are 
discussed.

12. What if I have important information for FAA and the comment period 
                               is closed?

    You may always provide FAA with written information after the close 
of the comment period and it will be considered if time permits. Because 
contacts after the close of the comment may not be seen by other 
interested persons, if they substantially and specifically influence the 
FAA's decision, we may need to reopen the comment period.



PART 13_INVESTIGATIVE AND ENFORCEMENT PROCEDURES--Table of Contents



                   Subpart A_Investigative Procedures

Sec.
13.1 Reports of violations.
13.3 Investigations (general).
13.5 Formal complaints.
13.7 Records, documents and reports.

                    Subpart B_Administrative Actions

13.11 Administrative disposition of certain violations.

                   Subpart C_Legal Enforcement Actions

13.13 Consent orders.
13.14 Civil penalties: General.
13.15 Civil penalties: Other than by administrative assessment.
13.16 Civil penalties: Administrative assessment against a person other 
          than an individual acting as a pilot, flight engineer, 
          mechanic, or repairman. Administrative assessment against all 
          persons for hazardous materials violations.
13.17 Seizure of aircraft.
13.18 Civil penalties: Administrative assessment against an individual 
          acting as a pilot, flight engineer, mechanic, or repairman.
13.19 Certificate action.
13.20 Orders of compliance, cease and desist orders, orders of denial, 
          and other orders.
13.21 Military personnel.
13.23 Criminal penalties.
13.25 Injunctions.
13.27 Final order of Hearing Officer in certificate of aircraft 
          registration proceedings.
13.29 Civil penalties: Streamlined enforcement procedures for certain 
          security violations.

              Subpart D_Rules of Practice for FAA Hearings

13.31 Applicability.
13.33 Appearances.
13.35 Request for hearing.
13.37 Hearing Officer's powers.
13.39 Disqualification of Hearing Officer.
13.41 [Reserved]
13.43 Service and filing of pleadings, motions, and documents.
13.44 Computation of time and extension of time.
13.45 Amendment of notice and answer.
13.47 Withdrawal of notice or request for hearing.
13.49 Motions.
13.51 Intervention.
13.53 Depositions.
13.55 Notice of hearing.
13.57 Subpoenas and witness fees.
13.59 Evidence.
13.61 Argument and submittals.
13.63 Record.

      Subpart E_Orders of Compliance Under the Hazardous Materials 
                           Transportation Act

13.71 Applicability.
13.73 Notice of proposed order of compliance.
13.75 Reply or request for hearing.
13.77 Consent order of compliance.
13.79 Hearing.
13.81 Order of immediate compliance.
13.83 Appeal.
13.85 Filing, service and computation of time.

[[Page 34]]

13.87 Extension of time.

     Subpart F_Formal Fact-Finding Investigation Under an Order of 
                              Investigation

13.101 Applicability.
13.103 Order of investigation.
13.105 Notification.
13.107 Designation of additional parties.
13.109 Convening the investigation.
13.111 Subpoenas.
13.113 Noncompliance with the investigative process.
13.115 Public proceedings.
13.117 Conduct of investigative proceeding or deposition.
13.119 Rights of persons against self-incrimination.
13.121 Witness fees.
13.123 Submission by party to the investigation.
13.125 Depositions.
13.127 Reports, decisions and orders.
13.129 Post-investigation action.
13.131 Other procedures.

        Subpart G_Rules of Practice in FAA Civil Penalty Actions

13.201 Applicability.
13.202 Definitions.
13.203 Separation of functions.
13.204 Appearances and rights of parties.
13.205 Administrative law judges.
13.206 Intervention.
13.207 Certification of documents.
13.208 Complaint.
13.209 Answer.
13.210 Filing of documents.
13.211 Service of documents.
13.212 Computation of time.
13.213 Extension of time.
13.214 Amendment of pleadings.
13.215 Withdrawal of complaint or request for hearing.
13.216 Waivers.
13.217 Joint procedural or discovery schedule.
13.218 Motions.
13.219 Interlocutory appeals.
13.220 Discovery.
13.221 Notice of hearing.
13.222 Evidence.
13.223 Standard of proof.
13.224 Burden of proof.
13.225 Offer of proof.
13.226 Public disclosure of evidence.
13.227 Expert or opinion witnesses.
13.228 Subpoenas.
13.229 Witness fees.
13.230 Record.
13.231 Argument before the administrative law judge.
13.232 Initial decision.
13.233 Appeal from initial decision.
13.234 Petition to reconsider or modify a final decision and order of 
          the FAA decisionmaker on appeal.
13.235 Judicial review of a final decision and order.

          Subpart H_Civil Monetary Penalty Inflation Adjustment

13.301 Scope and purpose.
13.303 Definitions.
13.305 Cost of living adjustments of civil monetary penalties.

         Subpart I_Flight Operational Quality Assurance Programs

13.401 Flight Operational Quality Assurance program: Prohibition against 
          use of data for enforcement purposes.

    Authority: 18 U.S.C. 6002; 28 U.S.C. 2461 (note); 49 U.S.C. 106(g), 
5121-5128, 40113-40114, 44103-44106, 44701-44703, 44709-44710, 44713, 
46101-46111, 46301, 46302 (for a violation of 49 U.S.C. 46504), 46304-
46316, 46318, 46501-46502, 46504-46507, 47106, 47107, 47111, 47122, 
47306, 47531-47532; 49 CFR 1.47.

    Source: Docket No. 18884, 44 FR 63723, Nov. 5, 1979, unless 
otherwise noted.



                   Subpart A_Investigative Procedures



Sec. 13.1  Reports of violations.

    (a) Any person who knows of a violation of the Federal Aviation Act 
of 1958, as amended, the Hazardous Materials Transportation Act relating 
to the transportation or shipment by air of hazardous materials, the 
Airport and Airway Development Act of 1970, the Airport and Airway 
Improvement Act of 1982, the Airport and Airway Improvement Act of 1982 
as amended by the Airport and Airway Safety and Capacity Expansion Act 
of 1987, or any rule, regulation, or order issued thereunder, should 
report it to appropriate personnel of any FAA regional or district 
office.
    (b) Each report made under this section, together with any other 
information the FAA may have that is relevant to the matter reported, 
will be reviewed by FAA personnel to determine the nature and type of 
any additional investigation or enforcement action the FAA will take.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-17, 
53 FR 33783, Aug. 31, 1988]

[[Page 35]]



Sec. 13.3  Investigations (general).

    (a) Under the Federal Aviation Act of 1958, as amended, (49 U.S.C. 
1301 et seq.), the Hazardous Materials Transportation Act (49 U.S.C. 
1801 et seq.), the Airport and Airway Development Act of 1970 (49 U.S.C. 
1701 et seq.), the Airport and Airway Improvement Act of 1982 (49 U.S.C. 
2201 et seq.), the Airport and Airway Improvement Act of 1982 (as 
amended, 49 U.S.C. App. 2201 et seq., Airport and Airway Safety and 
Capacity Expansion Act of 1987), and the Regulations of the Office of 
the Secretary of Transportation (49 CFR 1 et seq.), the Administrator 
may conduct investigations, hold hearings, issue subpoenas, require the 
production of relevant documents, records, and property, and take 
evidence and depositions.
    (b) For the purpose of investigating alleged violations of the 
Federal Aviation Act of 1958, as amended the Hazardous Materials 
Transportation Act, the Airport and Airway Development Act of 1970, the 
Airport and Airway Improvement Act of 1982, the Airport and Airway 
Improvement Act of 1982 as amended by the Airport and Airway Safety and 
Capacity Expansion Act of 1987, or any rule, regulation, or order issued 
thereunder, the Administrator's authority has been delegated to the 
various services and or offices for matters within their respective 
areas for all routine investigations. When the compulsory processes of 
sections 313 and 1004 (49 U.S.C. 1354 and 1484) of the Federal Aviation 
Act, or section 109 of the Hazardous Materials Transportation Act (49 
U.S.C. 1808) are invoked, the Administrator's authority has been 
delegated to the Chief Counsel, the Deputy Chief Counsel, each Assistant 
Chief Counsel, each Regional Counsel, the Aeronautical Center Counsel, 
and the Technical Center Counsel.
    (c) In conducting formal investigations, the Chief Counsel, the 
Deputy Chief Counsel, each Assistant Chief Counsel, each Regional 
Counsel, the Aeronautical Center Counsel, and the Technical Center 
Counsel may issue an order of investigation in accordance with subpart F 
of this part.
    (d) A complaint against the sponsor, proprietor, or operator of a 
Federally-assisted airport involving violations of the legal authorities 
listed in Sec. 16.1 of this chapter shall be filed in accordance with 
the provisions of part 16 of this chapter, except in the case of 
complaints, investigations, and proceedings initiated before December 
16, 1996, the effective date of part 16 of this chapter.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-17, 
53 FR 33783, Aug. 31, 1988; 53 FR 35255, Sept. 12, 1988; Amdt. 13-19, 54 
FR 39290, Sept. 25, 1989; Amdt. 13-27, 61 FR 54004, Oct. 16, 1996; Amdt. 
13-29, 62 FR 46865, Sept. 4, 1997]



Sec. 13.5  Formal complaints.

    (a) Any person may file a complaint with the Administrator with 
respect to anything done or omitted to be done by any person in 
contravention of any provision of any Act or of any regulation or order 
issued under it, as to matters within the jurisdiction of the 
Administrator. This section does not apply to complaints against the 
Administrator or employees of the FAA acting within the scope of their 
employment.
    (b) Complaints filed under this section must--
    (1) Be submitted in writing and identified as a complaint filed for 
the purpose of seeking an appropriate order or other enforcement action;
    (2) Be submitted to the Federal Aviation Administration, Office of 
the Chief Counsel, Attention: Enforcement Docket (AGC-10), 800 
Independence Avenue, S.W., Washington, DC 20591;
    (3) Set forth the name and address, if known, of each person who is 
the subject of the complaint and, with respect to each person, the 
specific provisions of the Act or regulation or order that the 
complainant believes were violated;
    (4) Contain a concise but complete statement of the facts relied 
upon to substantiate each allegation;
    (5) State the name, address and telephone number of the person 
filing the complaint; and
    (6) Be signed by the person filing the complaint or a duly 
authorized representative.
    (c) Complaints which do not meet the requirements of paragraph (b) 
of this section will be considered reports under Sec. 13.1.

[[Page 36]]

    (d) Complaints which meet the requirements of paragraph (b) of this 
section will be docketed and a copy mailed to each person named in the 
complaint.
    (e) Any complaint filed against a member of the Armed Forces of the 
United States acting in the performance of official duties shall be 
referred to the Secretary of the Department concerned for action in 
accordance with the procedures set forth in Sec. 13.21 of this part.
    (f) The person named in the complaint shall file an answer within 20 
days after service of a copy of the complaint.
    (g) After the complaint has been answered or after the allotted time 
in which to file an answer has expired, the Administrator shall 
determine if there are reasonable grounds for investigating the 
complaint.
    (h) If the Administrator determines that a complaint does not state 
facts which warrant an investigation or action, the complaint may be 
dismissed without a hearing and the reason for the dismissal shall be 
given, in writing, to the person who filed the complaint and the person 
named in the complaint.
    (i) If the Administrator determines that reasonable grounds exist, 
an informal investigation may be initiated or an order of investigation 
may be issued in accordance with subpart F of this part, or both. Each 
person named in the complaint shall be advised which official has been 
delegated the responsibility under Sec. 13.3(b) or (c) for conducting 
the investigation.
    (j) If the investigation substantiates the allegations set forth in 
the complaint, a notice of proposed order may be issued or other 
enforcement action taken in accordance with this part.
    (k) The complaint and other pleadings and official FAA records 
relating to the disposition of the complaint are maintained in current 
docket form in the Enforcement Docket (AGC-10), Office of the Chief 
Counsel, Federal Aviation Administration, 800 Independence Avenue, S.W., 
Washington, D. C. 20591. Any interested person may examine any docketed 
material at that office, at any time after the docket is established, 
except material that is ordered withheld from the public under 
applicable law or regulations, and may obtain a photostatic or duplicate 
copy upon paying the cost of the copy.

(Secs. 313(a), 314(a), 601 through 610, and 1102 of the Federal Aviation 
Act of 1958 (49 U.S.C. 1354(a), 1421 through 1430, 1502); sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 13-14, 44 FR 63723, Nov. 5, 1979; as amended by Amdt. 13-16, 
45 FR 35307, May 27, 1980; Amdt. 13-19, 54 FR 39290, Sept. 25, 1989]



Sec. 13.7  Records, documents and reports.

    Each record, document and report that the Federal Aviation 
Regulations require to be maintained, exhibited or submitted to the 
Administrator may be used in any investigation conducted by the 
Administrator; and, except to the extent the use may be specifically 
limited or prohibited by the section which imposes the requirement, the 
records, documents and reports may be used in any civil penalty action, 
certificate action, or other legal proceeding.



                    Subpart B_Administrative Actions



Sec. 13.11  Administrative disposition of certain violations.

    (a) If it is determined that a violation or an alleged violation of 
the Federal Aviation Act of 1958, or an order or regulation issued under 
it, or of the Hazardous Materials Transportation Act, or an order or 
regulation issued under it, does not require legal enforcement action, 
an appropriate official of the FAA field office responsible for 
processing the enforcement case or other appropriate FAA official may 
take administrative action in disposition of the case.
    (b) An administrative action under this section does not constitute 
a formal adjudication of the matter, and may be taken by issuing the 
alleged violator--
    (1) A ``Warning Notice'' which recites available facts and 
information about the incident or condition and indicates that it may 
have been a violation; or
    (2) A ``Letter of Correction'' which confirms the FAA decision in 
the matter and states the necessary corrective action the alleged 
violator has taken

[[Page 37]]

or agrees to take. If the agreed corrective action is not fully 
completed, legal enforcement action may be taken.



                   Subpart C_Legal Enforcement Actions



Sec. 13.13  Consent orders.

    (a) At any time before the issuance of an order under this subpart, 
the official who issued the notice and the person subject to the notice 
may agree to dispose of the case by the issuance of a consent order by 
the official.
    (b) A proposal for a consent order, submitted to the official who 
issued the notice, under this section must include--
    (1) A proposed order;
    (2) An admission of all jurisdictional facts;
    (3) An express waiver of the right to further procedural steps and 
of all rights to judicial review; and
    (4) An incorporation by reference of the notice and an 
acknowledgment that the notice may be used to construe the terms of the 
order.
    (c) If the issuance of a consent order has been agreed upon after 
the filing of a request for hearing in accordance with subpart D of this 
part, the proposal for a consent order shall include a request to be 
filed with the Hearing Officer withdrawing the request for a hearing and 
requesting that the case be dismissed.



Sec. 13.14  Civil penalties: General.

    (a) Any person who violates any of the following statutory 
provisions, or any rule, regulation, or order issued thereunder, is 
subject to a civil penalty of not more than the amount specified in 49 
U.S.C. chapter 463 for each violation:
    (1) Chapter 401 (except sections 40103(a) and (d), 40105, 40116, and 
40117);
    (2) Chapter 441 (except section 44109);
    (3) Section 44502(b) or (c);
    (4) Chapter 447 (except sections 44717 and 44719-44723);
    (5) Chapter 451;
    (6) Sections 46301(b), 46302 (for a violation of 49 U.S.C. 46504), 
or 46318;
    (7) Section 47107(b); or
    (8) Sections 47528 through 47530.
    (b) Any person who knowingly commits an act in violation of 49 
U.S.C. chapter 51 or a regulation prescribed or order issued under that 
chapter, is subject to a civil penalty under 49 U.S.C. 5123.
    (c) The minimum and maximum amounts of civil penalties for 
violations of the statutory provisions specified in paragraphs (a) and 
(b) of this section, or rules, regulations, or orders issued thereunder, 
are periodically adjusted for inflation in accordance with the formula 
established in 28 U.S.C. 2461 note and implemented in 14 CFR part 13, 
subpart H.

[Amdt. 13-32; 69 FR 59495, Oct. 4, 2004; Amdt. 13-32; 70 FR 1813, Jan. 
11, 2005; 71 FR 70464, Dec. 5, 2006]



Sec. 13.15  Civil penalties: Other than by administrative assessment.

    (a) The FAA uses the procedures in this section when it seeks a 
civil penalty other than by the administrative assessment procedures in 
Sec. Sec. 13.16 or 13.18.
    (b) The authority of the Administrator, under 49 U.S.C. chapter 463, 
to seek a civil penalty for a violation cited in Sec. 13.14(a), and the 
ability to refer cases to the United States Attorney General, or the 
delegate of the Attorney General, for prosecution of civil penalty 
actions sought by the Administrator is delegated to the Chief Counsel; 
the Deputy Chief Counsel for Operations; the Assistant Chief Counsel for 
Enforcement; the Assistant Chief Counsel, Europe, Africa, and Middle 
East Area Office; the Regional Counsel; the Aeronautical Center Counsel; 
and the Technical Center Counsel. This delegation applies to cases 
involving:
    (1) An amount in controversy in excess of:
    (i) $50,000, if the violation was committed by any person before 
December 12, 2003;
    (ii) $400,000, if the violation was committed by a person other than 
an individual or small business concern on or after December 12, 2003;
    (iii) $50,000, if the violation was committed by an individual or 
small business concern on or after December 12, 2003; or
    (2) An in rem action, seizure of aircraft subject to lien, suit for 
injunctive relief, or for collection of an assessed civil penalty.

[[Page 38]]

    (c) The Administrator may compromise any civil penalty proposed 
under this section, before referral to the United States Attorney 
General, or the delegate of the Attorney General, for prosecution.
    (1) The Administrator, through the Chief Counsel; the Deputy Chief 
Counsel for Operations; the Assistant Chief Counsel for Enforcement; the 
Assistant Chief Counsel, Europe, Africa, and Middle East Area Office; 
the Regional Counsel; the Aeronautical Center Counsel; or the Technical 
Center Counsel sends a civil penalty letter to the person charged with a 
violation cited in Sec. 13.14(a). The civil penalty letter contains a 
statement of the charges, the applicable law, rule, regulation, or 
order, the amount of civil penalty that the Administrator will accept in 
full settlement of the action or an offer to compromise the civil 
penalty.
    (2) Not later than 30 days after receipt of the civil penalty 
letter, the person charged with a violation may present any material or 
information in answer to the charges to the agency attorney, either 
orally or in writing, that may explain, mitigate, or deny the violation 
or that may show extenuating circumstances. The Administrator will 
consider any material or information submitted in accordance with this 
paragraph to determine whether the person is subject to a civil penalty 
or to determine the amount for which the Administrator will compromise 
the action.
    (3) If the person charged with the violation offers to compromise 
for a specific amount, that person must send to the agency attorney a 
certified check or money order for that amount, payable to the Federal 
Aviation Administration. The Chief Counsel; the Deputy Chief Counsel for 
Operations; the Assistant Chief Counsel for Enforcement; the Assistant 
Chief Counsel, Europe, Africa, and Middle East Area Office; the Regional 
Counsel; Aeronautical Center Counsel; or the Technical Center Counsel 
may accept the certified check or money order or may refuse and return 
the certified check or money order.
    (4) If the offer to compromise is accepted by the Administrator, the 
agency attorney will send a letter to the person charged with the 
violation stating that the certified check or money order is accepted in 
full settlement of the civil penalty action.
    (5) If the parties cannot agree to compromise the civil penalty 
action or the offer to compromise is rejected and the certified check or 
money order submitted in compromise is returned, the Administrator may 
refer the civil penalty action to the United States Attorney General, or 
the delegate of the Attorney General, to begin proceedings in a United 
States district court, pursuant to the authority in 49 U.S.C. 46305, to 
prosecute and collect the civil penalty.

[Amdt. 13-18, 53 FR 34653, Sept. 7, 1988, as amended by Amdt. 13-20, 55 
FR 15128, Apr. 20, 1990; Amdt. 13-29, 62 FR 46865, Sept. 4, 1997; Amdt. 
13-32; 69 FR 59495, Oct. 4, 2004]



Sec. 13.16  Civil Penalties: Administrative assessment against a person
other than an individual acting as a pilot, flight engineer, mechanic,
or repairman. Administrative assessment against all persons for hazardous 
          materials violations.

    (a) The FAA uses these procedures when it assesses a civil penalty 
against a person other than an individual acting as a pilot, flight 
engineer, mechanic, or repairman for a violation cited in the first 
sentence of 49 U.S.C. 46301(d)(2) or in 49 U.S.C. 47531, or any 
implementing rule, regulation or order.
    (b) District court jurisdiction. Notwithstanding the provisions of 
paragraph (a) of this section, the United States district courts have 
exclusive jurisdiction of any civil penalty action initiated by the FAA 
for violations described in those paragraphs, under 49 U.S.C. 
46301(d)(4), if--
    (1) The amount in controversy is more than $50,000 for a violation 
committed by any person before December 12, 2003;
    (2) The amount in controversy is more than $400,000 for a violation 
committed by a person other than an individual or small business concern 
on or after December 12, 2003;
    (3) The amount in controversy is more than $50,000 for a violation 
committed by an individual or a small business concern on or after 
December 12, 2003;

[[Page 39]]

    (4) The action is in rem or another action in rem based on the same 
violation has been brought;
    (5) The action involves an aircraft subject to a lien that has been 
seized by the Government; or
    (6) Another action has been brought for an injunction based on the 
same violation.
    (c) Hazardous materials violations. The FAA may assess a civil 
penalty against any person who knowingly commits an act in violation of 
49 U.S.C. chapter 51 or a regulation prescribed or order issued under 
that chapter, under 49 U.S.C. 5123 and 49 CFR 1.47(k). An order 
assessing a civil penalty for a violation under 49 U.S.C. chapter 51, or 
a rule, regulation, or order issued thereunder, is issued only after the 
following factors have been considered:
    (1) The nature, circumstances, extent, and gravity of the violation;
    (2) With respect to the violator, the degree of culpability, any 
history of prior violations, the ability to pay, and any effect on the 
ability to continue to do business; and
    (3) Such other matters as justice may require.
    (d) Order assessing civil penalty. An order assessing civil penalty 
may be issued for a violation described in paragraphs (a) or (c) of this 
section, or as otherwise provided by statute, after notice and 
opportunity for a hearing. A person charged with a violation may be 
subject to an order assessing civil penalty in the following 
circumstances:
    (1) An order assessing civil penalty may be issued if a person 
charged with a violation submits or agrees to submit a civil penalty for 
a violation.
    (2) An order assessing civil penalty may be issued if a person 
charged with a violation does not request a hearing under paragraph 
(g)(2)(ii) of this section within 15 days after receipt of a final 
notice of proposed civil penalty.
    (3) Unless an appeal is filed with the FAA decisionmaker in a timely 
manner, an initial decision or order of an administrative law judge 
shall be considered an order assessing civil penalty if an 
administrative law judge finds that an alleged violation occurred and 
determines that a civil penalty, in an amount found appropriate by the 
administrative law judge, is warranted.
    (4) Unless a petition for review is filed with a U.S. Court of 
Appeals in a timely manner, a final decision and order of the 
Administrator shall be considered an order assessing civil penalty if 
the FAA decisionmaker finds that an alleged violation occurred and a 
civil penalty is warranted.
    (e) Delegation of authority. (1) The authority of the Administrator 
under 49 U.S.C. 46301(d), 47531, and 5123, and 49 CFR 1.47(k) to 
initiate and assess civil penalties for a violation of those statutes or 
a rule, regulation, or order issued thereunder, is delegated to the 
Deputy Chief Counsel for Operations; the Assistant Chief Counsel for 
Enforcement; the Assistant Chief Counsel, Europe, Africa, and Middle 
East Area Office; the Regional Counsel; the Aeronautical Center Counsel; 
and the Technical Center Counsel.
    (2) The authority of the Administrator under 49 U.S.C. 5123, 49 CFR 
1.47(k), 49 U.S.C. 46301(d), and 49 U.S.C. 46305 to refer cases to the 
Attorney General of the United States, or the delegate of the Attorney 
General, for collection of civil penalties is delegated to the Deputy 
Chief Counsel for Operations; the Assistant Chief Counsel for 
Enforcement; Assistant Chief Counsel, Europe, Africa, and Middle East 
Area Office; the Regional Counsel; the Aeronautical Center Counsel; and 
the Technical Center Counsel.
    (3) The authority of the Administrator under 49 U.S.C. 46301(f) to 
compromise the amount of a civil penalty imposed is delegated to the 
Deputy Chief Counsel for Operations; the Assistant Chief Counsel for 
Enforcement; Assistant Chief Counsel, Europe, Africa, and Middle East 
Area Office; the Regional Counsel; the Aeronautical Center Counsel; and 
the Technical Center Counsel.
    (4) The authority of the Administrator under 49 U.S.C. 5123 (e) and 
(f) and 49 CFR 1.47(k) to compromise the amount of a civil penalty 
imposed is delegated to the Deputy Chief Counsel for Operations; the 
Assistant Chief Counsel for Enforcement; Assistant Chief Counsel, 
Europe, Africa, and Middle East Area Office; the Regional Counsel; the 
Aeronautical Center Counsel; and the Technical Center Counsel.

[[Page 40]]

    (f) Notice of proposed civil penalty. A civil penalty action is 
initiated by sending a notice of proposed civil penalty to the person 
charged with a violation or to the agent for services for the person 
under 49 U.S.C. 46103. A notice of proposed civil penalty will be sent 
to the individual charged with a violation or to the president of the 
corporation or company charged with a violation. In response to a notice 
of proposed civil penalty, a corporation or company may designate in 
writing another person to receive documents in that civil penalty 
action. The notice of proposed civil penalty contains a statement of the 
charges and the amount of the proposed civil penalty. Not later than 30 
days after receipt of the notice of proposed civil penalty, the person 
charged with a violation shall--
    (1) Submit the amount of the proposed civil penalty or an agreed-
upon amount, in which case either an order assessing civil penalty or 
compromise order shall be issued in that amount;
    (2) Submit to the agency attorney one of the following:
    (i) Written information, including documents and witness statements, 
demonstrating that a violation of the regulations did not occur or that 
a penalty or the amount of the penalty is not warranted by the 
circumstances.
    (ii) A written request to reduce the proposed civil penalty, the 
amount of reduction, and the reasons and any documents supporting a 
reduction of the proposed civil penalty, including records indicating a 
financial inability to pay or records showing that payment of the 
proposed civil penalty would prevent the person from continuing in 
business.
    (iii) A written request for an informal conference to discuss the 
matter with the agency attorney and to submit relevant information or 
documents; or
    (3) Request a hearing, in which case a complaint shall be filed with 
the hearing docket clerk.
    (g) Final notice of proposed civil penalty. A final notice of 
proposed civil penalty may be issued after participation in informal 
procedures provided in paragraph (f)(2) of this section or failure to 
respond in a timely manner to a notice of proposed civil penalty. A 
final notice of proposed civil penalty will be sent to the individual 
charged with a violation, to the president of the corporation or company 
charged with a violation, or a person previously designated in writing 
by the individual, corporation, or company to receive documents in that 
civil penalty action. If not previously done in response to a notice of 
proposed civil penalty, a corporation or company may designate in 
writing another person to receive documents in that civil penalty 
action. The final notice of proposed civil penalty contains a statement 
of the charges and the amount of the proposed civil penalty and, as a 
result of information submitted to the agency attorney during informal 
procedures, may modify an allegation or a proposed civil penalty 
contained in a notice of proposed civil penalty.
    (1) A final notice of proposed civil penalty may be issued--
    (i) If the person charged with a violation fails to respond to the 
notice of proposed civil penalty within 30 days after receipt of that 
notice; or
    (ii) If the parties participated in any informal procedures under 
paragraph (f)(2) of this section and the parties have not agreed to 
compromise the action or the agency attorney has not agreed to withdraw 
the notice of proposed civil penalty.
    (2) Not later than 15 days after receipt of the final notice of 
proposed civil penalty, the person charged with a violation shall do one 
of the following--
    (i) Submit the amount of the proposed civil penalty or an agreed-
upon amount, in which case either an order assessing civil penalty or a 
compromise order shall be issued in that amount; or
    (ii) Request a hearing, in which case a complaint shall be filed 
with the hearing docket clerk.
    (h) Request for a hearing. Any person charged with a violation may 
request a hearing, pursuant to paragraph (f)(3) or paragraph (g)(2)(ii) 
of this section, to be conducted in accordance with the procedures in 
subpart G of this part. A person requesting a hearing shall file a 
written request for a hearing with the hearing docket clerk, using the 
appropriate address set forth in Sec. 13.210(a) of

[[Page 41]]

this part, and shall mail a copy of the request to the agency attorney. 
The request for a hearing may be in the form of a letter but must be 
dated and signed by the person requesting a hearing. The request for a 
hearing may be typewritten or may be legibly handwritten.
    (i) Hearing. If the person charged with a violation requests a 
hearing pursuant to paragraph (f)(3) or paragraph (g)(2)(ii) of this 
section, the original complaint shall be filed with the hearing docket 
clerk and a copy shall be sent to the person requesting the hearing. The 
procedural rules in subpart G of this part apply to the hearing and any 
appeal. At the close of the hearing, the administrative law judge shall 
issue, either orally on the record or in writing, an initial decision, 
including the reasons for the decision, that contains findings or 
conclusions on the allegations contained, and the civil penalty sought, 
in the complaint.
    (j) Appeal. Either party may appeal the administrative law judge's 
initial decision to the FAA decisionmaker pursuant to the procedures in 
subpart G of this part. If a party files a notice of appeal pursuant to 
Sec. 13.233 of subpart G, the effectiveness of the initial decision is 
stayed until a final decision and order of the Administrator have been 
entered on the record. The FAA decisionmaker shall review the record and 
issue a final decision and order of the Administrator that affirm, 
modify, or reverse the initial decision. The FAA decisionmaker may 
assess a civil penalty but shall not assess a civil penalty in an amount 
greater than that sought in the complaint.
    (k) Payment. A person shall pay a civil penalty by sending a 
certified check or money order, payable to the Federal Aviation 
Administration, to the agency attorney.
    (l) Collection of civil penalties. If an individual does not pay a 
civil penalty imposed by an order assessing civil penalty or other final 
order, the Administrator may take action provided under the law to 
collect the penalty.
    (m) A party may seek review only of a final decision and order of 
the FAA decisionmaker involving a violation of the Federal aviation 
statute or the Federal hazardous materials transportation law. Judicial 
review is in the United States Court of Appeals for the District of 
Columbia Circuit or the United States court of appeals for the circuit 
in which the party resides or has the party's principal place of 
business as provided in Sec. 13.235 of this part. Neither an initial 
decision or an order issued by an administrative law judge that has not 
been appealed to the FAA decisionmaker, nor an order compromising a 
civil penalty action, may be appealed under any of those sections.
    (n) Compromise. The FAA may compromise the amount of any civil 
penalty imposed under this section, under 49 U.S.C. 5123(e), 46031(f), 
46303(b), or 46318 at any time before referring the action to the United 
States Attorney General, or the delegate of the Attorney General, for 
collection.
    (1) An agency attorney may compromise any civil penalty action where 
a person charged with a violation agrees to pay a civil penalty and the 
FAA agrees not to make a finding of violation. Under such agreement, a 
compromise order is issued following the payment of the agreed-on amount 
or the signing of a promissory note. The compromise order states the 
following:
    (i) The person has paid a civil penalty or has signed a promissory 
note providing for installment payments.
    (ii) The FAA makes no finding of a violation.
    (iii) The compromise order shall not be used as evidence of a prior 
violation in any subsequent civil penalty proceeding or certificate 
action proceeding.
    (2) An agency attorney may compromise the amount of a civil penalty 
proposed in a notice, assessed in an order, or imposed in a compromise 
order.

[Amdt. 13-32; 70 FR 1813, Jan. 11, 2005; 70 FR 2925, Jan. 18, 2005, as 
amended at 70 FR 8238, Feb. 18, 2005; 71 FR 70464, Dec. 5, 2006]



Sec. 13.17  Seizure of aircraft.

    (a) Under section 903 of the Federal Aviation Act of 1958 (49 U.S.C. 
1473), a State or Federal law enforcement officer, or a Federal Aviation 
Administration safety inspector, authorized in an order of seizure 
issued by the Regional

[[Page 42]]

Administrator of the region, or by the Chief Counsel, may summarily 
seize an aircraft that is involved in a violation for which a civil 
penalty may be imposed on its owner or operator.
    (b) Each person seizing an aircraft under this section shall place 
it in the nearest available and adequate public storage facility in the 
judicial district in which it was seized.
    (c) The Regional Administrator or Chief Counsel, without delay, 
sends a written notice and a copy of this section, to the registered 
owner of the seized aircraft, and to each other persons shown by FAA 
records to have an interest in it, stating the--
    (1) Time, date, and place of seizure;
    (2) Name and address of the custodian of the aircraft;
    (3) Reasons for the seizure, including the violations believed, or 
judicially determined, to have been committed; and
    (4) Amount that may be tendered as--
    (i) A compromise of a civil penalty for the alleged violation; or
    (ii) Payment for a civil penalty imposed by a Federal court for a 
proven violation.
    (d) The Chief Counsel, or the Regional Counsel or Assistant Chief 
Counsel for the region or area in which an aircraft is seized under this 
section, immediately sends a report to the United States District 
Attorney for the judicial district in which it was seized, requesting 
the District Attorney to institute proceedings to enforce a lien against 
the aircraft.
    (e) The Regional Administrator or Chief Counsel directs the release 
of a seized aircraft whenever--
    (1) The alleged violator pays a civil penalty or an amount agreed 
upon in compromise, and the costs of seizing, storing, and maintaining 
the aircraft;
    (2) The aircraft is seized under an order of a Federal Court in 
proceedings in rem to enforce a lien against the aircraft, or the United 
States District Attorney for the judicial district concerned notifies 
the FAA that the District Attorney refuses to institute those 
proceedings; or
    (3) A bond in the amount and with the sureties prescribed by the 
Chief Counsel, the Regional Counsel, or the Assistant Chief Counsel is 
deposited, conditioned on payment of the penalty, or the compromise 
amount, and the costs of seizing, storing, and maintaining the aircraft.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-19, 
54 FR 39290, Sept. 25, 1989; Amdt. 13-29, 62 FR 46865, Sept. 4, 1997]



Sec. 13.18  Civil penalties: Administrative assessment against an 
individual acting as a pilot, flight engineer, mechanic, or repairman.

    (a) General. (1) This section applies to each action in which the 
FAA seeks to assess a civil penalty by administrative procedures against 
an individual acting as a pilot, flight engineer, mechanic, or 
repairman, under 49 U.S.C. 46301(d)(5), for a violation listed in 49 
U.S.C. 46301(d)(2). This section does not apply to a civil penalty 
assessed for violation of 49 U.S.C. chapter 51, or a rule, regulation, 
or order issued thereunder.
    (2) District court jurisdiction. Notwithstanding the provisions of 
paragraph (a)(1) of this section, the United States district courts have 
exclusive jurisdiction of any civil penalty action involving an 
individual acting as a pilot, flight engineer, mechanic, or repairman 
for violations described in that paragraph, under 49 U.S.C. 46301(d)(4), 
if:
    (i) The amount in controversy is more than $50,000.
    (ii) The action involves an aircraft subject to a lien that has been 
seized by the Government; or
    (iii) Another action has been brought for an injunction based on the 
same violation.
    (b) Definitions. As used in this part, the following definitions 
apply:
    (1) Flight engineer means an individual who holds a flight engineer 
certificate issued under part 63 of this chapter.
    (2) Individual acting as a pilot, flight engineer, mechanic, or 
repairman means an individual acting in such capacity, whether or not 
that individual holds the respective airman certificate issued by the 
FAA.
    (3) Mechanic means an individual who holds a mechanic certificate 
issued under part 65 of this chapter.

[[Page 43]]

    (4) Pilot means an individual who holds a pilot certificate issued 
under part 61 of this chapter.
    (5) Repairman means an individual who holds a repairman certificate 
issued under part 65 of this chapter.
    (c) Delegation of authority. (1) The authority of the Administrator 
under 49 U.S.C. 46301(d)(5), to initiate and assess civil penalties is 
delegated to the Chief Counsel; the Deputy Chief Counsel for Operations; 
the Assistant Chief Counsel for Enforcement; Assistant Chief Counsel, 
Europe, Africa, and Middle East Area Office; the Regional Counsel; the 
Aeronautical Center Counsel; and the Technical Center Counsel.
    (2) The authority of the Administrator to refer cases to the 
Attorney General of the United States, or the delegate of the Attorney 
General, for collection of civil penalties is delegated to the Chief 
Counsel; the Deputy Chief Counsel for Operations; the Assistant Chief 
Counsel for Enforcement; Assistant Chief Counsel, Europe, Africa, and 
Middle East Area Office; the Regional Counsel; the Aeronautical Center 
Counsel; and the Technical Center Counsel.
    (3) The authority of the Administrator to compromise the amount of a 
civil penalty under 49 U.S.C. 46301(f) is delegated to the Chief 
Counsel; the Deputy Chief Counsel for Operations; the Assistant Chief 
Counsel for Enforcement; Assistant Chief Counsel, Europe, Africa, and 
Middle East Area Office; the Regional Counsel; the Aeronautical Center 
Counsel; and the Technical Center Counsel.
    (d) Notice of proposed assessment. A civil penalty action is 
initiated by sending a notice of proposed assessment to the individual 
charged with a violation specified in paragraph (a) of this section. The 
notice of proposed assessment contains a statement of the charges and 
the amount of the proposed civil penalty. The individual charged with a 
violation may do the following:
    (1) Submit the amount of the proposed civil penalty or an agreed-on 
amount, in which case either an order of assessment or a compromise 
order will be issued in that amount.
    (2) Answer the charges in writing.
    (3) Submit a written request for an informal conference to discuss 
the matter with an agency attorney and submit relevant information or 
documents.
    (4) Request that an order be issued in accordance with the notice of 
proposed assessment so that the individual charged may appeal to the 
National Transportation Safety Board.
    (e) Failure to respond to notice of proposed assessment. An order of 
assessment may be issued if the individual charged with a violation 
fails to respond to the notice of proposed assessment within 15 days 
after receipt of that notice.
    (f) Order of assessment. An order of assessment, which assesses a 
civil penalty, may be issued for a violation described in paragraph (a) 
of this section after notice and an opportunity to answer any charges 
and be heard as to why such order should not be issued.
    (g) Appeal. Any individual who receives an order of assessment 
issued under this section may appeal the order to the National 
Transportation Safety Board. The appeal stays the effectiveness of the 
Administrator's order.
    (h) Exhaustion of administrative remedies. An individual 
substantially affected by an order of the NTSB or the Administrator may 
petition for review only of a final decision and order of the National 
Transportation Safety Board to a court of appeals of the United States 
for the circuit in which the individual charged resides or has his or 
her principal place of business or the United States Court of Appeals 
for the District of Columbia Circuit, under 49 U.S.C. 46110 and 
46301(d)(6). Neither an order of assessment that has not been appealed 
to the National Transportation Board, nor an order compromising a civil 
penalty action, may be appealed under those sections.
    (i) Compromise. The FAA may compromise any civil penalty action 
initiated under this section, in accordance with 49 U.S.C. 46301(f).
    (1) An agency attorney may compromise any civil penalty action where 
an individual charged with a violation agrees to pay a civil penalty and 
the

[[Page 44]]

FAA agrees to make no finding of violation. Under such agreement, a 
compromise order is issued following the payment of the agreed-on amount 
or the signing of a promissory note. The compromise order states the 
following:
    (i) The individual has paid a civil penalty or has signed a 
promissory note providing for installment payments;
    (ii) The FAA makes no finding of violation; and
    (iii) The compromise order will not be used as evidence of a prior 
violation in any subsequent civil penalty proceeding or certificate 
action proceeding.
    (2) An agency attorney may compromise the amount of any civil 
penalty proposed or assessed in an order.
    (j) Payment. (1) An individual must pay a civil penalty by:
    (i) Sending a certified check or money order, payable to the Federal 
Aviation Administration, to the FAA office identified in the order of 
assessment, or
    (ii) Making an electronic funds transfer according to the directions 
specified in the order of assessment.
    (2) The civil penalty must be paid within 30 days after service of 
the order of assessment, unless an appeal is filed with the National 
Transportation Safety Board. The civil penalty must be paid within 30 
days after a final order of the Board or the Court of Appeals affirms 
the order of assessment in whole or in part.
    (k) Collection of civil penalties. If an individual does not pay a 
civil penalty imposed by an order of assessment or other final order, 
the Administrator may take action provided under the law to collect the 
penalty.

[Amdt. 13-32; 69 FR 59497, Oct. 4, 2004]



Sec. 13.19  Certificate action.

    (a) Under section 609 of the Federal Aviation Act of 1958 (49 U.S.C. 
1429), the Administrator may reinspect any civil aircraft, aircraft 
engine, propeller, appliance, air navigation facility, or air agency, 
and may re-examine any civil airman. Under section 501(e) of the FA Act, 
any Certificate of Aircraft Registration may be suspended or revoked by 
the Administrator for any cause that renders the aircraft ineligible for 
registration.
    (b) If, as a result of such a reinspection re-examination, or other 
investigation made by the Administrator under section 609 of the FA Act, 
the Administrator determines that the public interest and safety in air 
commerce requires it, the Administrator may issue an order amending, 
suspending, or revoking, all or part of any type certificate, production 
certificate, airworthiness certificate, airman certificate, air carrier 
operating certificate, air navigation facility certificate, or air 
agency certificate. This authority may be exercised for remedial 
purposes in cases involving the Hazardous Materials Transportation Act 
(49 U.S.C. 1801 et seq.) or regulations issued under that Act. This 
authority is also exercised by the Chief Counsel, the Assistant Chief 
Counsel, Enforcement, the Assistant Chief Counsel, Regulations, the 
Assistant Chief Counsel, Europe, Africa, and Middle East Area Office, 
each Regional Counsel, and the Aeronautical Center Counsel. If the 
Administrator finds that any aircraft registered under Part 47 of this 
chapter is ineligible for registration, the Administrator issues an 
order suspending or revoking that certificate. This authority as to 
aircraft found ineligible for registration is also exercised by each 
Regional Counsel, the Aeronautical Center Counsel, and the Assistant 
Chief Counsel, Europe, Africa, and Middle East Area Office.
    (c) Before issuing an order under paragraph (b) of this section, the 
Chief Counsel, the Assistant Chief Counsel, Enforcement, the Assistant 
Chief Counsel, Regulations, the Assistant Chief Counsel, Europe, Africa, 
and Middle East Area Office, each Regional Counsel, or the Aeronautical 
Center Counsel advises the certificate holder of the charges or other 
reasons upon which the Administrator bases the proposed action and, 
except in an emergency, allows the holder to answer any charges and to 
be heard as to why the certificate should not be amended, suspended, or 
revoked. The holder may, by checking the appropriate box on the form 
that is sent to the holder with the notice of proposed certificate 
action, elect to--

[[Page 45]]

    (1) Admit the charges and surrender his or her certificate;
    (2) Answer the charges in writing;
    (3) Request that an order be issued in accordance with the notice of 
proposed certificate action so that the certificate holder may appeal to 
the National Transportation Safety Board, if the charges concerning a 
matter under Title VI of the FA Act;
    (4) Request an opportunity to be heard in an informal conference 
with the FAA counsel; or
    (5) Request a hearing in accordance with Subpart D of this part if 
the charges concern a matter under Title V of the FA Act.

Except as provided in Sec. 13.35(b), unless the certificate holder 
returns the form and, where required, an answer or motion, with a 
postmark of not later than 15 days after the date of receipt of the 
notice, the order of the Administrator is issued as proposed. If the 
certificate holder has requested an informal conference with the FAA 
counsel and the charges concern a matter under Title V of the FA Act, 
the holder may after that conference also request a formal hearing in 
writing with a postmark of not later than 10 days after the close of the 
conference. After considering any information submitted by the 
certificate holder, the Chief Counsel, the Assistant Chief Counsel for 
Regulations and Enforcement, the Regional Counsel concerned, or the 
Aeronautical Center Counsel (as to matters under Title V of the FA Act) 
issues the order of the Administrator, except that if the holder has 
made a valid request for a formal hearing on a matter under Title V of 
the FA Act initially or after an informal conference, Subpart D of this 
part governs further proceedings.
    (d) Any person whose certificate is affected by an order issued 
under this section may appeal to the National Transportation Safety 
Board. If the certificate holder files an appeal with the Board, the 
Administrator's order is stayed unless the Administrator advises the 
Board that an emergency exists and safety in air commerce requires that 
the order become effective immediately. If the Board is so advised, the 
order remains effective and the Board shall finally dispose of the 
appeal within 60 days after the date of the advice. This paragraph does 
not apply to any person whose Certificate of Aircraft Registration is 
affected by an order issued under this section.

[Doc. No. 13-14, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-15, 
45 FR 20773, Mar. 31, 1980; Amdt. 13-19, 54 FR 39290, Sept. 25, 1989; 
Amdt. 13-29, 62 FR 46865, Sept. 4, 1997; 75 FR 41979, July 20, 2010]



Sec. 13.20  Orders of compliance, cease and desist orders, orders of
denial, and other orders.

    (a) This section applies to orders of compliance, cease and desist 
orders, orders of denial, and other orders issued by the Administrator 
to carry out the provisions of the Federal Aviation Act of 1958, as 
amended, the Hazardous Materials Transportation Act, the Airport and 
Airway Development Act of 1970, and the Airport and Airway Improvement 
Act of 1982, or the Airport and Airway Improvement Act of 1982 as 
amended by the Airport and Airway Safety and Capacity Expansion Act of 
1987. This section does not apply to orders issued pursuant to section 
602 or section 609 of the Federal Aviation Act of 1958, as amended.
    (b) Unless the Administrator determines that an emergency exists and 
safety in air commerce requires the immediate issuance of an order under 
this section, the person subject to the order shall be provided with 
notice prior to issuance.
    (c) Within 30 days after service of the notice, the person subject 
to the order may''
    (1) Request an opportunity to be heard in an informal conference 
with an FAA attorney;
    (2) Reply in writing; or
    (3) Request a hearing in accordance with subpart D of this part.
    (d) If an informal conference is held or a reply is filed, as to any 
charges not withdrawn or not subject to a consent order, the person 
subject to the order may, within 10 days after receipt of notice that 
the remaining charges are not withdrawn, request a hearing in accordance 
with subpart D of this part.
    (e) Failure to request a hearing within the period provided in 
paragraphs (c) or (d) of this section--

[[Page 46]]

    (1) Constitutes a waiver of the right to appeal and the right to a 
hearing, and
    (2) Authorizes the official who issued the notice to find the facts 
to be as alleged in the notice, or as modified as the official may 
determine necessary based on any written response, and to issue an 
appropriate order, without further notice or proceedings.
    (f) If a hearing is requested in accordance with paragraph (c) or 
(d) of this section, the procedure of Subpart D of this part applies. At 
the close of the hearing, the Hearing Officer, on the record or 
subsequently in writing, shall set forth findings and conclusions and 
the reasons therefor, and either--
    (1) Dismiss the notice; or
    (2) Issue an order.
    (g) Any party to the hearing may appeal from the order of the 
Hearing Officer by filing a notice of appeal with the Administrator 
within 20 days after the date of issuance of the order.
    (h) If a notice of appeal is not filed from the order issued by a 
Hearing Officer, such order is the final agency order.
    (i) Any person filing an appeal authorized by paragraph (g) of this 
section shall file an appeal brief with the Administrator within 40 days 
after the date of issuance of the order, and serve a copy on the other 
party. A reply brief must be filed within 20 days after service of the 
appeal brief and a copy served on the appellant.
    (j) On appeal the Administrator reviews the available record of the 
proceeding, and issues an order dismissing, reversing, modifying or 
affirming the order. The Administrator's order includes the reasons for 
the Administrator's action.
    (k) For good cause shown, requests for extensions of time to file 
any document under this section may be granted by--
    (1) The official who issued the order, if the request is filed prior 
to the designation of a Hearing Officer; or
    (2) The Hearing Officer, if the request is filed prior to the filing 
of a notice of appeal; or
    (3) The Administrator, if the request is filed after the filing of a 
notice of appeal.
    (l) Except in the case of an appeal from the decision of a Hearing 
Officer, the authority of the Administrator under this section is also 
exercised by the Chief Counsel, Deputy Chief Counsel, each Assistant 
Chief Counsel, each Regional Counsel, and the Aeronautical Center 
Counsel (as to matters under Title V of the Federal Aviation Act of 
1958).
    (m) Filing and service of documents under this section shall be 
accomplished in accordance with Sec. 13.43; and the periods of time 
specified in this section shall be computed in accordance with Sec. 
13.44.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-17, 
53 FR 33783, Aug. 31, 1988; Amdt. 13-19, 54 FR 39290, Sept. 25, 1989; 
Amdt. 13-29, 62 FR 46865, Sept. 4, 1997; Amdt. 13-36, 79 FR 46967, Aug. 
12, 2014]



Sec. 13.21  Military personnel.

    If a report made under this part indicates that, while performing 
official duties, a member of the Armed Forces, or a civilian employee of 
the Department of Defense who is subject to the Uniform Code of Military 
Justice (10 U.S.C. Ch. 47), has violated the Federal Aviation Act of 
1958, or a regulation or order issued under it, the Chief Counsel, the 
Assistant Chief Counsel, Enforcement, the Assistant Chief Counsel, 
Regulations, the Assistant Chief Counsel, Europe, Africa, and Middle 
East Area Office, each Regional Counsel, and the Aeronautical Center 
Counsel send a copy of the report to the appropriate military authority 
for such disciplinary action as that authority considers appropriate and 
a report to the Administrator thereon.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-19, 
54 FR 39290, Sept. 25, 1989; Amdt. 13-29, 62 FR 46866, Sept. 4, 1997]



Sec. 13.23  Criminal penalties.

    (a) Sections 902 and 1203 of the Federal Aviation Act of 1958 (49 
U.S.C. 1472 and 1523), provide criminal penalties for any person who 
knowingly and willfully violates specified provisions of that Act, or 
any regulation or order issued under those provisions. Section 110(b) of 
the Hazardous Materials Transportation Act (49 U.S.C. 1809(b)) provides 
for a criminal penalty of a

[[Page 47]]

fine of not more than $25,000, imprisonment for not more than five 
years, or both, for any person who willfully violates a provision of 
that Act or a regulation or order issued under it.
    (b) If an inspector or other employee of the FAA becomes aware of a 
possible violation of any criminal provision of the Federal Aviation Act 
of 1958 (except a violation of section 902 (i) through (m) which is 
reported directly to the Federal Bureau of Investigation), or of the 
Hazardous Materials Transportation Act, relating to the transportation 
or shipment by air of hazardous materials, he or she shall report it to 
the Office of the Chief Counsel or the Regional Counsel or Assistant 
Chief Counsel for the region or area concerned. If appropriate, that 
office refers the report to the Department of Justice for criminal 
prosecution of the offender. If such an inspector or other employee 
becomes aware of a possible violation of a Federal statute that is 
within the investigatory jurisdiction of another Federal agency, he or 
she shall immediately report it to that agency according to standard FAA 
practices.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-19, 
54 FR 39290, Sept. 25, 1989; Amdt. 13-29, 62 FR 46866, Sept. 4, 1997]



Sec. 13.25  Injunctions.

    (a) Whenever it is determined that a person has engaged, or is about 
to engage, in any act or practice constituting a violation of the 
Federal Aviation Act of 1958, or any regulation or order issued under it 
for which the FAA exercises enforcement responsibility, or, with respect 
to the transportation or shipment by air of any hazardous materials, in 
any act or practice constituting a violation of the Hazardous Materials 
Transportation Act, or any regulation or order issued under it for which 
the FAA exercises enforcement responsibility, the Chief Counsel, the 
Assistant Chief Counsel, Enforcement, the Assistant Chief Counsel, 
Regulations, the Assistant Chief Counsel, Europe, Africa, and Middle 
East Area Office, each Regional Counsel, and the Aeronautical Center 
Counsel may request the United States Attorney General, or the delegate 
of the Attorney General, to bring an action in the appropriate United 
States District Court for such relief as is necessary or appropriate, 
including mandatory or prohibitive injunctive relief, interim equitable 
relief, and punitive damages, as provided by section 1007 of the Federal 
Aviation Act of 1958 (49 U.S.C. 1487) and section 111(a) of the 
Hazardous Materials Transportation Act (49 U.S.C. 1810).
    (b) Whenever it is determined that there is substantial likelihood 
that death, serious illness, or severe personal injury, will result from 
the transportation by air of a particular hazardous material before an 
order of compliance proceeding, or other administrative hearing or 
formal proceeding to abate the risk of the harm can be completed, the 
Chief Counsel, the Assistant Chief Counsel, Enforcement, the Assistant 
Chief Counsel, Regulations, the Assistant Chief Counsel, Europe, Africa, 
and Middle East Area Office, each Regional Counsel, and the Aeronautical 
Center Counsel may bring, or request the United States Attorney General 
to bring, an action in the appropriate United States District Court for 
an order suspending or restricting the transportation by air of the 
hazardous material or for such other order as is necessary to eliminate 
or ameliorate the imminent hazard, as provided by section 111(b) of the 
Hazardous Materials Transportation Act (49 U.S.C. 1810).

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-19, 
54 FR 39290, Sept. 25, 1989; Amdt. 13-29, 62 FR 46866, Sept. 4, 1997]



Sec. 13.27  Final order of Hearing Officer in certificate of aircraft
registration proceedings.

    (a) If, in proceedings under section 501(b) of the Federal Aviation 
Act of 1958 (49 U.S.C. 1401), the Hearing Officer determines that the 
aircraft is ineligible for a Certificate of Aircraft Registration, the 
Hearing Officer shall suspend or revoke the respondent's certificate, as 
proposed in the notice of proposed certificate action.
    (b) If the final order of the Hearing Officer makes a decision on 
the merits, it shall contain a statement of the findings and conclusions 
of law on all material issues of fact and law. If the

[[Page 48]]

Hearing Officer finds that the allegations of the notice have been 
proven, but that no sanction is required, the Hearing Officer shall make 
appropriate findings and issue an order terminating the notice. If the 
Hearing Officer finds that the allegations of the notice have not been 
proven, the Hearing Officer shall issue an order dismissing the notice. 
If the Hearing Officer finds it to be equitable and in the public 
interest, the Hearing Officer shall issue an order terminating the 
proceeding upon payment by the respondent of a civil penalty in an 
amount agreed upon by the parties.
    (c) If the order is issued in writing, it shall be served upon the 
parties.

[Doc. No. 13-14, 44 FR 63723, Nov. 5, 1979; as amended by Amdt. 13-15, 
45 FR 20773, Mar. 31, 1980; Amdt. 13-34, 75 FR 41979, July 20, 2010]



Sec. 13.29  Civil penalties: Streamlined enforcement procedures for
certain security violations.

    This section may be used, at the agency's discretion, in enforcement 
actions involving individuals presenting dangerous or deadly weapons for 
screening at airports or in checked baggage where the amount of the 
proposed civil penalty is less than $5,000. In these cases, sections 
13.16(a), 13.16(c), and 13.16 (f) through (l) of this chapter are used, 
as well as paragraphs (a) through (d) of this section:
    (a) Delegation of authority. The authority of the Administrator, 
under 49 U.S.C. 46301, to initiate the assessment of civil penalties for 
a violation of 49 U.S.C. Subtitle VII, or a rule, regulation, or order 
issued thereunder, is delegated to the regional Civil Aviation Security 
Division Manager and the regional Civil Aviation Security Deputy 
Division Manager for the purpose of issuing notices of violation in 
cases involving violations of 49 U.S.C. Subtitle VII and the FAA's 
regulations by individuals presenting dangerous or deadly weapons for 
screening at airport checkpoints or in checked baggage. This authority 
may not be delegated below the level of the regional Civil Aviation 
Security Deputy Division Manager.
    (b) Notice of violation. A civil penalty action is initiated by 
sending a notice of violation to the person charged with the violation. 
The notice of violation contains a statement of the charges and the 
amount of the proposed civil penalty. Not later than 30 days after 
receipt of the notice of violation, the person charged with a violation 
shall:
    (1) Submit the amount of the proposed civil penalty or an agreed-
upon amount, in which case either an order assessing a civil penalty or 
a compromise order shall be issued in that amount; or
    (2) Submit to the agency attorney identified in the material 
accompanying the notice any of the following:
    (i) Written information, including documents and witness statements, 
demonstrating that a violation of the regulations did not occur or that 
a penalty or the penalty amount is not warranted by the circumstances; 
or
    (ii) A written request to reduce the proposed civil penalty, the 
amount of reduction, and the reasons and any documents supporting a 
reduction of the proposed civil penalty, including records indicating a 
financial inability to pay or records showing that payment of the 
proposed civil penalty would prevent the person from continuing in 
business; or
    (iii) A written request for an informal conference to discuss the 
matter with an agency attorney and submit relevant information or 
documents; or
    (3) Request a hearing in which case a complaint shall be filed with 
the hearing docket clerk.
    (c) Final notice of violation and civil penalty assessment order. A 
final notice of violation and civil penalty assessment order (``final 
notice and order'') may be issued after participation in any informal 
proceedings as provided in paragraph (b)(2) of this section, or after 
failure of the respondent to respond in a timely manner to a notice of 
violation. A final notice and order will be sent to the individual 
charged with a violation. The final notice and order will contain a 
statement of the charges and the amount of the proposed civil penalty 
and, as a result of information submitted to the agency attorney during 
any informal procedures, may reflect a modified allegation or proposed 
civil penalty.
    A final notice and order may be issued--

[[Page 49]]

    (1) If the person charged with a violation fails to respond to the 
notice of violation within 30 days after receipt of that notice; or
    (2) If the parties participated in any informal procedures under 
paragraph (b)(2) of this section and the parties have not agreed to 
compromise the action or the agency attorney has not agreed to withdraw 
the notice of violation.
    (d) Order assessing civil penalty. An order assessing civil penalty 
may be issued after notice and opportunity for a hearing. A person 
charged with a violation may be subject to an order assessing civil 
penalty in the following circumstances:
    (1) An order assessing civil penalty may be issued if a person 
charged with a violation submits, or agrees to submit, the amount of 
civil penalty proposed in the notice of violation.
    (2) An order assessing civil penalty may be issued if a person 
charged with a violation submits, or agrees to submit, an agreed-upon 
amount of civil penalty that is not reflected in either the notice of 
violation or the final notice and order.
    (3) The final notice and order becomes (and contains a statement so 
indicating) an order assessing a civil penalty when the person charged 
with a violation submits the amount of the proposed civil penalty that 
is reflected in the final notice and order.
    (4) The final notice and order becomes (and contains a statement so 
indicating) an order assessing a civil penalty 16 days after receipt of 
the final notice and order, unless not later than 15 days after receipt 
of the final notice and order, the person charged with a violation does 
one of the following--
    (i) Submits an agreed-upon amount of civil penalty that is not 
reflected in the final notice and order, in which case an order 
assessing civil penalty or a compromise order shall be issued in that 
amount; or
    (ii) Requests a hearing in which case a complaint shall be filed 
with the hearing docket clerk.
    (5) Unless an appeal is filed with the FAA decisionmaker in a timely 
manner, an initial decision or order of an administrative law judge 
shall be considered an order assessing civil penalty if an 
administrative law judge finds that an alleged violation occurred and 
determines that a civil penalty, in an amount found to be appropriate by 
the administrative law judge, is warranted.
    (6) Unless a petition for review is filed with a U.S. Court of 
Appeals in a timely manner, a final decision and order of the 
Administrator shall be considered an order assessing civil penalty if 
the FAA decisionmaker finds that an alleged violation occurred and a 
civil penalty is warranted.

[Doc. No. 27873, 61 FR 44155, Aug. 28, 1996]



              Subpart D_Rules of Practice for FAA Hearings



Sec. 13.31  Applicability.

    This subpart applies to proceedings in which a hearing has been 
requested in accordance with Sec. Sec. 13.19(c)(5), 13.20(c), 13.20(d), 
13.75(a)(2), 13.75(b), or 13.81(e).

[Amdt. 13-18, 53 FR 34655, Sept. 7, 1988]



Sec. 13.33  Appearances.

    Any party to a proceeding under this subpart may appear and be heard 
in person or by attorney.



Sec. 13.35  Request for hearing.

    (a) A request for hearing must be made in writing and filed in the 
Hearing Docket.
    (1) If delivery is in person, or by expedited courier service. A 
person delivering the request for hearing in person or sending the 
request for hearing by commercial expedited courier (for example, 
Federal Express or United Parcel Service), should use the following 
address: Federal Aviation Administration, 600 Independence Avenue, SW., 
Wilbur Wright Building--Suite 2W1000, Washington, DC 20591; Attention: 
Hearing Docket Clerk, AGC-430.
    (2) If delivery is by U.S. Mail: If the request for hearing is sent 
by U.S. Mail, then it should be addressed as follows: Federal Aviation 
Administration, 800 Independence Avenue, SW., Washington, DC 20591: 
Attention: Hearing Docket Clerk, AGC-430, Wilbur Wright Building--Suite 
2W1000.
    (b) The request for hearing must describe the action proposed by the 
FAA, and must contain a statement that a

[[Page 50]]

hearing is requested. A copy of the request for hearing and a copy of 
the answer required by paragraph (c) of this section must be served on 
the official who issued the notice of proposed action.
    (c) An answer to the notice of proposed action must be filed with 
the request for hearing. All allegations in the notice not specifically 
denied in the answer are deemed admitted.
    (d) Within 15 days after service of the copy of the request for 
hearing, the official who issued the notice of proposed action forwards 
a copy of that notice, which serves as the complaint, to the Hearing 
Docket.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-19, 
54 FR 39290, Sept. 25, 1989; 71 FR 70464, Dec. 5, 2006]



Sec. 13.37  Hearing Officer's powers.

    Any Hearing Officer may--
    (a) Give notice concerning, and hold, prehearing conferences and 
hearings;
    (b) Administrator oaths and affirmations;
    (c) Examine witnesses;
    (d) Adopt procedures for the submission of evidence in written form;
    (e) Issue subpoenas and take depositions or cause them to be taken;
    (f) Rule on offers of proof;
    (g) Receive evidence;
    (h) Regulate the course of the hearing;
    (i) Hold conferences, before and during the hearing, to settle and 
simplify issues by consent of the parties;
    (j) Dispose of procedural requests and similar matters; and
    (k) Issue decisions, make findings of fact, make assessments, and 
issue orders, as appropriate.



Sec. 13.39  Disqualification of Hearing Officer.

    If disqualified for any reason, the Hearing Officer shall withdraw 
from the case.



Sec. 13.41  [Reserved]



Sec. 13.43  Service and filing of pleadings, motions, and documents.

    (a) Copies of all pleadings, motions, and documents filed with the 
Hearing Docket must be served upon all parties to the proceedings by the 
person filing them.
    (b) Service may be made by personal delivery or by mail.
    (c) A certificate of service shall accompany all documents when they 
are tendered for filing and shall consist of a certificate of personal 
delivery or a certificate of mailing, executed by the person making the 
personal delivery or mailing the document.
    (d) Whenever proof of service by mail is made, the date of mailing 
or the date as shown on the postmark shall be the date of service, and 
where personal service is made, the date of personal delivery shall be 
the date of service.
    (e) The date of filing is the date the document is actually 
received.



Sec. 13.44  Computation of time and extension of time.

    (a) In computing any period of time prescribed or allowed by this 
subpart, the date of the act, event, default, notice or order after 
which the designated period of time begins to run is not to be included 
in the computation. The last day of the period so computed is to be 
included unless it is a Saturday, Sunday, or legal holiday for the FAA, 
in which event the period runs until the end of the next day which is 
neither a Saturday, Sunday nor a legal holiday.
    (b) Upon written request filed with the Hearing Docket and served 
upon all parties, and for good cause shown, a Hearing Officer may grant 
an extension of time to file any documents specified in this subpart.



Sec. 13.45  Amendment of notice and answer.

    At any time more than 10 days before the date of hearing, any party 
may amend his or her notice, answer, or other pleading, by filing the 
amendment with the Hearing Officer and serving a copy of it on each 
other party. After that time, amendments may be allowed only in the 
discretion of the Hearing Officer. If an amendment to an initial 
pleading has been allowed, the Hearing Officer shall allow the other 
parties a reasonable opportunity to answer.

[[Page 51]]



Sec. 13.47  Withdrawal of notice or request for hearing.

    At any time before the hearing, the FAA counsel may withdraw the 
notice of proposed action, and the party requesting the hearing may 
withdraw the request for hearing.



Sec. 13.49  Motions.

    (a) Motion to dismiss for insufficiency. A respondent who requests a 
formal hearing may, in place of an answer, file a motion to dismiss for 
failure of the allegations in the notice of proposed action to state a 
violation of the FA Act or of this chapter or to show lack of 
qualification of the respondent. If the Hearing Officer denies the 
motion, the respondent shall file an answer within 10 days.
    (b) [Reserved]
    (c) Motion for more definite statement. The certificate holder may, 
in place of an answer, file a motion that the allegations in the notice 
be made more definite and certain. If the Hearing Officer grants the 
motion, the FAA counsel shall comply within 10 days after the date it is 
granted. If the Hearing Officer denies the motion the certificate holder 
shall file an answer within 10 days after the date it is denied.
    (d) Motion for judgment on the pleadings. After the pleadings are 
closed, either party may move for a judgment on the pleadings.
    (e) Motion to strike. Upon motion of either party, the Hearing 
Officer may order stricken, from any pleadings, any insufficient 
allegation or defense, or any immaterial, impertinent, or scandalous 
matter.
    (f) Motion for production of documents. Upon motion of any party 
showing good cause, the Hearing Officer may, in the manner provided by 
Rule 34, Federal Rules of Civil Procedure, order any party to produce 
any designated document, paper, book, account, letter, photograph, 
object, or other tangible thing, that is not privileged, that 
constitutes or contains evidence relevant to the subject matter of the 
hearings, and that is in the party's possession, custody, or control.
    (g) Consolidation of motions. A party who makes a motion under this 
section shall join with it all other motions that are then available to 
the party. Any objection that is not so raised is considered to be 
waived.
    (h) Answers to motions. Any party may file an answer to any motion 
under this section within 5 days after service of the motion.



Sec. 13.51  Intervention.

    Any person may move for leave to intervene in a proceeding and may 
become a party thereto, if the Hearing Officer, after the case is sent 
to the Hearing Officer for hearing, finds that the person may be bound 
by the order to be issued in the proceedings or has a property or 
financial interest that may not be adequately represented by existing 
parties, and that the intervention will not unduly broaden the issues or 
delay the proceedings. Except for good cause shown, a motion for leave 
to intervene may not be considered if it is filed less than 10 days 
before the hearing.



Sec. 13.53  Depositions.

    After the respondent has filed a request for hearing and an answer, 
either party may take testimony by deposition in accordance with section 
1004 of the Federal Aviation Act of 1958 (49 U.S.C. 1484) or Rule 26, 
Federal Rules of Civil Procedure.



Sec. 13.55  Notice of hearing.

    The Hearing Officer shall set a reasonable date, time, and place for 
the hearing, and shall give the parties adequate notice thereof and of 
the nature of the hearing. Due regard shall be given to the convenience 
of the parties with respect to the place of the hearing.



Sec. 13.57  Subpoenas and witness fees.

    (a) The Hearing Officer to whom a case is assigned may, upon 
application by any party to the proceeding, issue subpoenas requiring 
the attendance of witnesses or the production of documentary or tangible 
evidence at a hearing or for the purpose of taking depositions. However, 
the application for producing evidence must show its general relevance 
and reasonable scope. This paragraph does not apply to the attendance of 
FAA employees or to the production of documentary evidence in

[[Page 52]]

the custody of such an employee at a hearing.
    (b) A person who applies for the production of a document in the 
custody of an FAA employee must follow the procedure in Sec. 13.49(f). 
A person who applies for the attendance of an FAA employee must send the 
application, in writing, to the Hearing Officer setting forth the need 
for that employee's attendance.
    (c) A witness in a proceeding under this subpart is entitled to the 
same fees and mileage as is paid to a witness in a court of the United 
States under comparable circumstances. The party at whose instance the 
witness is subpoenaed or appears shall pay the witness fees.
    (d) Notwithstanding the provisions of paragraph (c) of this section, 
the FAA pays the witness fees and mileage if the Hearing Officer who 
issued the subpoena determines, on the basis of a written request and 
good cause shown, that--
    (1) The presence of the witness will materially advance the 
proceeding; and
    (2) The party at whose instance the witness is subpoenaed would 
suffer a serious hardship if required to pay the witness fees and 
mileage.



Sec. 13.59  Evidence.

    (a) Each party to a hearing may present the party's case or defense 
by oral or documentary evidence, submit evidence in rebuttal, and 
conduct such cross-examination as may be needed for a full disclosure of 
the facts.
    (b) Except with respect to affirmative defenses and orders of 
denial, the burden of proof is upon the FAA counsel.
    (c) The Hearing Officer may order information contained in any 
report or document filed or in any testimony given pursuant to this 
subpart withheld from public disclosure when, in the judgment of the 
Hearing Officer, disclosure would adversely affect the interests of any 
person and is not required in the public interest or is not otherwise 
required by statute to be made available to the public. Any person may 
make written objection to the public disclosure of such information, 
stating the ground for such objection.



Sec. 13.61  Argument and submittals.

    The Hearing Officer shall give the parties adequate opportunity to 
present arguments in support of motions, objections, and the final 
order. The Hearing Officer may determine whether arguments are to be 
oral or written. At the end of the hearing the Hearing Officer may, in 
the discretion of the Hearing Officer, allow each party to submit 
written proposed findings and conclusions and supporting reasons for 
them.



Sec. 13.63  Record.

    The testimony and exhibits presented at a hearing, together with all 
papers, requests, and rulings filed in the proceedings are the exclusive 
basis for the issuance of an order. Either party may obtain a transcript 
from the official reporter upon payment of the fees fixed therefor.



      Subpart E_Orders of Compliance Under the Hazardous Materials 
                           Transportation Act



Sec. 13.71  Applicability.

    Whenever the Chief Counsel, the Assistant Chief Counsel, 
Enforcement, the Assistant Chief Counsel, Europe, Africa, and Middle 
East Area Office, or a Regional Counsel has reason to believe that a 
person is engaging in the transportation or shipment by air of hazardous 
materials in violation of the Hazardous Materials Transportation Act, or 
any regulation or order issued under it for which the FAA exercises 
enforcement responsibility, and the circumstances do not require the 
issuance of an order of immediate compliance, he may conduct proceedings 
pursuant to section 109 of that Act (49 U.S.C. 1808) to determine the 
nature and extent of the violation, and may thereafter issue an order 
directing compliance.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-19, 
54 FR 39290, Sept. 25, 1989; Amdt. 13-29, 62 FR 46866, Sept. 4, 1997]

[[Page 53]]



Sec. 13.73  Notice of proposed order of compliance.

    A compliance order proceeding commences when the Chief Counsel, the 
Assistant Chief Counsel, Enforcement, the Assistant Chief Counsel, 
Europe, Africa, and Middle East Area Office, or a Regional Counsel sends 
the alleged violator a notice of proposed order of compliance advising 
the alleged violator of the charges and setting forth the remedial 
action sought in the form of a proposed order of compliance.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-19, 
54 FR 39290, Sept. 25, 1989; Amdt. 13-29, 62 FR 46866, Sept. 4, 1997]



Sec. 13.75  Reply or request for hearing.

    (a) Within 30 days after service upon the alleged violator of a 
notice of proposed order of compliance, the alleged violator may--
    (1) File a reply in writing with the official who issued the notice; 
or
    (2) Request a hearing in accordance with subpart D of this part.
    (b) If a reply is filed, as to any charges not dismissed or not 
subject to a consent order of compliance, the alleged violator may, 
within 10 days after receipt of notice that the remaining charges are 
not dismissed, request a hearing in accordance with subpart D of this 
part.
    (c) Failure of the alleged violator to file a reply or request a 
hearing within the period provided in paragraph (a) or (b) of this 
section--
    (1) Constitutes a waiver of the right to a hearing and the right to 
an appeal, and
    (2) Authorizes the official who issued the notice to find the facts 
to be as alleged in the notice and to issue an appropriate order 
directing compliance, without further notice or proceedings.



Sec. 13.77  Consent order of compliance.

    (a) At any time before the issuance of an order of compliance, the 
official who issued the notice and the alleged violator may agree to 
dispose of the case by the issuance of a consent order of compliance by 
the official.
    (b) A proposal for a consent order submitted to the official who 
issued the notice under this section must include--
    (1) A proposed order of compliance;
    (2) An admission of all jurisdictional facts;
    (3) An express waiver of right to further procedural steps and of 
all rights to judicial review;
    (4) An incorporation by reference of the notice and an 
acknowledgement that the notice may be used to construe the terms of the 
order of compliance; and
    (5) If the issuance of a consent order has been agreed upon after 
the filing of a request for hearing in accordance with subpart D of this 
part, the proposal for a consent order shall include a request to be 
filed with the Hearing Officer withdrawing the request for a hearing and 
requesting that the case be dismissed.



Sec. 13.79  Hearing.

    If an alleged violator requests a hearing in accordance with Sec. 
13.75, the procedure of subpart D of this part applies. At the close of 
the hearing, the Hearing Officer, on the record or subsequently in 
writing, sets forth the Hearing Officer's findings and conclusion and 
the reasons therefor, and either--
    (a) Dismisses the notice of proposed order of compliance; or
    (b) Issues an order of compliance.



Sec. 13.81  Order of immediate compliance.

    (a) Notwithstanding Sec. Sec. 13.73 through 13.79, the Chief 
Counsel, the Assistant Chief Counsel, Enforcement, the Assistant Chief 
Counsel, Europe, Africa, and Middle East Area Office, or a Regional 
Counsel may issue an order of immediate compliance, which is effective 
upon issuance, if the person who issues the order finds that--
    (1) There is strong probability that a violation is occurring or is 
about to occur;
    (2) The violation poses a substantial risk to health or to safety of 
life or property; and
    (3) The public interest requires the avoidance or amelioration of 
that risk through immediate compliance and waiver of the procedures 
afforded under Sec. Sec. 13.73 through 13.79.
    (b) An order of immediate compliance is served promptly upon the 
person against whom the order is issued

[[Page 54]]

by telephone or telegram, and a written statement of the relevant facts 
and the legal basis for the order, including the findings required by 
paragraph (a) of this section, is served promptly by personal service or 
by mail.
    (c) The official who issued the order of immediate compliance may 
rescind or suspend the order if it appears that the criteria set forth 
in paragraph (a) of this section are no longer satisfied, and, when 
appropriate, may issue a notice of proposed order of compliance under 
Sec. 13.73 in lieu thereof.
    (d) If at any time in the course of a proceeding commenced in 
accordance with Sec. 13.73 the criteria set forth in paragraph (a) of 
this section are satisfied, the official who issued the notice may issue 
an order of immediate compliance, even if the period for filing a reply 
or requesting a hearing specified in Sec. 13.75 has not expired.
    (e) Within three days after receipt of service of an order of 
immediate compliance, the alleged violator may request a hearing in 
accordance with Subpart D of this part and the procedure in that subpart 
will apply except that--
    (1) The case will be heard within fifteen days after the date of the 
order of immediate compliance unless the alleged violator requests a 
later date;
    (2) The order will serve as the complaint; and
    (3) The Hearing Officer shall issue his decision and order 
dismissing, reversing, modifying, or affirming the order of immediate 
compliance on the record at the close of the hearing.
    (f) The filing of a request for hearing in accordance with paragraph 
(e) of this section does not stay the effectiveness of an order of 
immediate compliance.
    (g) At any time after an order of immediate compliance has become 
effective, the official who issued the order may request the United 
States Attorney General, or the delegate of the Attorney General, to 
bring an action for appropriate relief in accordance with Sec. 13.25.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 13-19, 
54 FR 39290, Sept. 25, 1989; Amdt. 13-29, 62 FR 46866, Sept. 4, 1997]



Sec. 13.83  Appeal.

    (a) Any party to the hearing may appeal from the order of the 
Hearing Officer by filing a notice of appeal with the Administrator 
within 20 days after the date of issuance of the order.
    (b) Any person against whom an order of immediate compliance has 
been issued in accordance with Sec. 13.81 or the official who issued 
the order of immediate compliance may appeal from the order of the 
Hearing Officer by filing a notice of appeal with the Administrator 
within three days after the date of issuance of the order by the Hearing 
Officer.
    (c) Unless the Administrator expressly so provides, the filing of a 
notice of appeal does not stay the effectiveness of an order of 
immediate compliance.
    (d) If a notice of appeal is not filed from the order of compliance 
issued by a Hearing Officer, such order is the final agency order of 
compliance.
    (e) Any person filing an appeal authorized by paragraph (a) of this 
section shall file an appeal brief with the Administrator within 40 days 
after the date of the issuance of the order, and serve a copy on the 
other party. Any reply brief must be filed within 20 days after service 
of the appeal brief. A copy of the reply brief must be served on the 
appellant.
    (f) Any person filing an appeal authorized by paragraph (b) of this 
section shall file an appeal brief with the Administrator with the 
notice of appeal and serve a copy on the other party. Any reply brief 
must be filed within 3 days after receipt of the appeal brief. A copy of 
the reply brief must be served on the appellant.
    (g) On appeal the Administrator reviews the available record of the 
proceeding, and issues an order dismissing, reversing, modifying or 
affirming the order of compliance or the order of immediate compliance. 
The Administrator's order includes the reasons for the action.
    (h) In cases involving an order of immediate compliance, the 
Administrator's order on appeal is issued within ten days after the 
filing of the notice of appeal.

[[Page 55]]



Sec. 13.85  Filing, service and computation of time.

    Filing and service of documents under this subpart shall be 
accomplished in accordance with Sec. 13.43 except service of orders of 
immediate compliance under Sec. 13.81(b); and the periods of time 
specified in this subpart shall be computed in accordance with Sec. 
13.44.



Sec. 13.87  Extension of time.

    (a) The official who issued the notice of proposed order of 
compliance, for good cause shown, may grant an extension of time to file 
any document specified in this subpart, except documents to be filed 
with the Administrator.
    (b) Extensions of time to file documents with the Administrator may 
be granted by the Administrator upon written request, served upon all 
parties, and for good cause shown.



     Subpart F_Formal Fact-Finding Investigation Under an Order of 
                              Investigation



Sec. 13.101  Applicability.

    (a) This subpart applies to fact-finding investigations in which an 
order of investigation has been issued under Sec. 13.3(c) or Sec. 
13.5(i) of this part.
    (b) This subpart does not limit the authority of duly designated 
persons to issue subpoenas, administer oaths, examine witnesses and 
receive evidence in any informal investigation as provided for in 
sections 313 and 1004(a) of the Federal Aviation Act (49 U.S.C. 1354 and 
1484(a)) and section 109(a) of the Hazardous Materials Transportation 
Act (49 U.S.C. 1808(a)).



Sec. 13.103  Order of investigation.

    The order of investigation--
    (a) Defines the scope of the investigation by describing the 
information sought in terms of its subject matter or its relevancy to 
specified FAA functions;
    (b) Sets forth the form of the investigation which may be either by 
individual deposition or investigative proceeding or both; and
    (c) Names the official who is authorized to conduct the 
investigation and serve as the Presiding Officer.



Sec. 13.105  Notification.

    Any person under investigation and any person required to testify 
and produce documentary or physical evidence during the investigation 
will be advised of the purpose of the investigation, and of the place 
where the investigative proceeding or deposition will be convened. This 
may be accomplished by a notice of investigation or by a subpoena. A 
copy of the order of investigation may be sent to such persons, when 
appropriate.



Sec. 13.107  Designation of additional parties.

    (a) The Presiding Officer may designate additional persons as 
parties to the investigation, if in the discretion of the Presiding 
Officer, it will aid in the conduct of the investigation.
    (b) The Presiding Officer may designate any person as a party to the 
investigation if that person--
    (1) Petitions the Presiding Officer to participate as a party; and
    (2) Is so situated that the disposition of the investigation may as 
a practical matter impair the ability to protect that person's interest 
unless allowed to participate as a party, and
    (3) Is not adequately represented by existing parties.



Sec. 13.109  Convening the investigation.

    The investigation shall be conducted at such place or places 
designated by the Presiding Officer, and as convenient to the parties 
involved as expeditious and efficient handling of the investigation 
permits.



Sec. 13.111  Subpoenas.

    (a) Upon motion of the Presiding Officer, or upon the request of a 
party to the investigation, the Presiding Officer may issue a subpoena 
directing any person to appear at a designated time and place to testify 
or to produce documentary or physical evidence relating to any matter 
under investigation.
    (b) Subpoenas shall be served by personal service, or upon an agent 
designated in writing for the purpose, or by registered or certified 
mail addressed to such person or agent. Whenever service is made by 
registered or certified mail, the date of mailing shall

[[Page 56]]

be considered as the time when service is made.
    (c) Subpoenas shall extend in jurisdiction throughout the United 
States or any territory or possession thereof.



Sec. 13.113  Noncompliance with the investigative process.

    If any person fails to comply with the provisions of this subpart or 
with any subpoena or order issued by the Presiding Officer or the 
designee of the Presiding Officer, judicial enforcement may be initiated 
against that person under applicable statutes.



Sec. 13.115  Public proceedings.

    (a) All investigative proceedings and depositions shall be public 
unless the Presiding Officer determines that the public interest 
requires otherwise.
    (b) The Presiding Officer may order information contained in any 
report or document filed or in any testimony given pursuant to this 
subpart withheld from public disclosure when, in the judgment of the 
Presiding Officer, disclosure would adversely affect the interests of 
any person and is not required in the public interest or is not 
otherwise required by statute to be made available to the public. Any 
person may make written objection to the public disclosure of such 
information, stating the grounds for such objection.



Sec. 13.117  Conduct of investigative proceeding or deposition.

    (a) The Presiding Officer or the designee of the Presiding Officer 
may question witnesses.
    (b) Any witness may be accompanied by counsel.
    (c) Any party may be accompanied by counsel and either the party or 
counsel may--
    (1) Question witnesses, provided the questions are relevant and 
material to the matters under investigation and would not unduly impede 
the progress of the investigation; and
    (2) Make objections on the record and argue the basis for such 
objections.
    (d) Copies of all notices or written communications sent to a party 
or witness shall upon request be sent to that person's attorney of 
record.



Sec. 13.119  Rights of persons against self-incrimination.

    (a) Whenever a person refuses, on the basis of a privilege against 
self-incrimination, to testify or provide other information during the 
course of any investigation conducted under this subpart, the Presiding 
Officer may, with the approval of the Attorney General of the United 
States, issue an order requiring the person to give testimony or provide 
other information. However, no testimony or other information so 
compelled (or any information directly or indirectly derived from such 
testimony or other information) may be used against the person in any 
criminal case, except in a prosecution for perjury, giving a false 
statement, or otherwise failing to comply with the order.
    (b) The Presiding Officer may issue an order under this section if--
    (1) The testimony or other information from the witness may be 
necessary to the public interest; and
    (2) The witness has refused or is likely to refuse to testify or 
provide other information on the basis of a privilege against self-
incrimination.
    (c) Immunity provided by this section will not become effective 
until the person has refused to testify or provide other information on 
the basis of a privilege against self-incrimination, and an order under 
this section has been issued. An order, however, may be issued 
prospectively to become effective in the event of a claim of the 
privilege.



Sec. 13.121  Witness fees.

    All witnesses appearing shall be compensated at the same rate as a 
witness appearing before a United States District Court.



Sec. 13.123  Submission by party to the investigation.

    (a) During an investigation conducted under this subpart, a party 
may submit to the Presiding Officer--
    (1) A list of witnesses to be called, specifying the subject matter 
of the expected testimony of each witness, and
    (2) A list of exhibits to be considered for inclusion in the record.
    (b) If the Presiding Officer determines that the testimony of a 
witness

[[Page 57]]

or the receipt of an exhibit in accordance with paragraph (a) of this 
section will be relevant, competent and material to the investigation, 
the Presiding Officer may subpoena the witness or use the exhibit during 
the investigation.



Sec. 13.125  Depositions.

    Depositions for investigative purposes may be taken at the 
discretion of the Presiding Officer with reasonable notice to the party 
under investigation. Such depositions shall be taken before the 
Presiding Officer or other person authorized to administer oaths and 
designated by the Presiding Officer. The testimony shall be reduced to 
writing by the person taking the deposition, or under the direction of 
that person, and where possible shall then be subscribed by the 
deponent. Any person may be compelled to appear and testify and to 
produce physical and documentary evidence.



Sec. 13.127  Reports, decisions and orders.

    The Presiding Officer shall issue a written report based on the 
record developed during the formal investigation, including a summary of 
principal conclusions. A summary of principal conclusions shall be 
prepared by the official who issued the order of investigation in every 
case which results in no action, or no action as to a particular party 
to the investigation. All such reports shall be furnished to the parties 
to the investigation and filed in the public docket. Insertion of the 
report in the Public Docket shall constitute ``entering of record'' and 
publication as prescribed by section 313(b) of the Federal Aviation Act.



Sec. 13.129  Post-investigation action.

    A decision on whether to initiate subsequent action shall be made on 
the basis of the record developed during the formal investigation and 
any other information in the possession of the Administrator.



Sec. 13.131  Other procedures.

    Any question concerning the scope or conduct of a formal 
investigation not covered in this subpart may be ruled on by the 
Presiding Officer on motion of the Presiding Officer, or on the motion 
of a party or a person testifying or producing evidence.



        Subpart G_Rules of Practice in FAA Civil Penalty Actions

    Source: Amdt. 13-21, 55 FR 27575, July 3, 1990, unless otherwise 
noted.



Sec. 13.201  Applicability.

    (a) This subpart applies to all civil penalty actions initiated 
under Sec. 13.16 of this part in which a hearing has been requested.
    (b) This subpart applies only to proceedings initiated after 
September 7, 1988. All other cases, hearings, or other proceedings 
pending or in progress before September 7, 1988, are not affected by the 
rules in this subpart.

[Amdt. 13-21, 55 FR 27575, July 3, 1990, as amended by Amdt. 13-32; 69 
FR 59497, Oct. 4, 2004]



Sec. 13.202  Definitions.

    Administrative law judge means an administrative law judge appointed 
pursuant to the provisions of 5 U.S.C. 3105.
    Agency attorney means the Deputy Chief Counsel for Operations, the 
Assistant Chief Counsel, Enforcement, the Assistant Chief Counsel, 
Europe, Africa, and Middle East Area Office, each Regional Counsel, the 
Aeronautical Center Counsel, or the Technical Center Counsel, or an 
attorney on the staff of the Assistant Chief Counsel, Enforcement, the 
Assistant Chief Counsel, Europe, Africa, and Middle East Area Office, 
each Regional Counsel, the Aeronautical Center Counsel, or the Technical 
Center Counsel who prosecutes a civil penalty action. An agency attorney 
shall not include:
    (1) The Chief Counsel, the Deputy Chief Counsel for Policy and 
Adjudication, or the Assistant Chief Counsel for Litigation;
    (2) Any attorney on the staff of the Assistant Chief Counsel for 
Litigation;
    (3) Any attorney who is supervised in a civil penalty action by a 
person who provides such advice to the FAA decisionmaker in that action 
or a factually-related action.
    Attorney means a person licensed by a state, the District of 
Columbia, or a

[[Page 58]]

territory of the United States to practice law or appear before the 
courts of that state or territory.
    Complaint means a document issued by an agency attorney alleging a 
violation of a provision of the Federal aviation statute listed in the 
first sentence of 49 U.S.C. 46301(d)(2) or in 49 U.S.C. 47531, or of the 
Federal hazardous materials transportation statute, 49 U.S.C. 5121-5128, 
or a rule, regulation, or order issued under those statutes, that has 
been filed with the Hearing Docket after a hearing has been requested 
under Sec. 13.16(f)(3) or (g)(2)(ii) of this part.
    FAA decisionmaker means the Administrator of the Federal Aviation 
Administration, acting in the capacity of the decisionmaker on appeal, 
or any person to whom the Administrator has delegated the 
Administrator's decisionmaking authority in a civil penalty action. As 
used in this subpart, the FAA decisionmaker is the official authorized 
to issue a final decision and order of the Administrator in a civil 
penalty action.
    Mail includes U.S. certified mail, U.S. registered mail, or use of 
an overnight express courier service.
    Order assessing civil penalty means a document that contains a 
finding of a violation of a provision of the Federal aviation statute 
listed in the first sentence of 49 U.S.C. 46301(d)(2) or in 49 U.S.C. 
47531, or of the Federal hazardous materials transportation statute, 49 
U.S.C. 5121-5128, or a rule, regulation or order issued under those 
statutes, and may direct payment of a civil penalty. Unless an appeal is 
filed with the FAA decisionmaker in a timely manner, an initial decision 
or order of an administrative law judge shall be considered an order 
assessing civil penalty if an administrative law judge finds that an 
alleged violation occurred and determines that a civil penalty, in an 
amount found appropriate by the administrative law judge, is warranted. 
Unless a petition for review is filed with a U.S. Court of Appeals in a 
timely manner, a final decision and order of the Administrator shall be 
considered an order assessing civil penalty if the FAA decisionmaker 
finds that an alleged violation occurred and a civil penalty is 
warranted.
    Party means the respondent or the Federal Aviation Administration 
(FAA).
    Personal delivery includes hand-delivery or use of a contract or 
express messenger service. ``Personal delivery'' does not include the 
use of Government interoffice mail service.
    Pleading means a complaint, an answer, and any amendment of these 
documents permitted under this subpart.
    Properly addressed means a document that shows an address contained 
in agency records, a residential, business, or other address submitted 
by a person on any document provided under this subpart, or any other 
address shown by other reasonable and available means.
    Respondent means a person, corporation, or company named in a 
complaint.

[Amdt. 13-21, 55 FR 27575, July 3, 1990, as amended by Amdt. 13-24, 58 
FR 50241, Sept. 24, 1993; Amdt. 13-29, 62 FR 46866, Sept. 4, 1997; 70 FR 
8238, Feb. 18, 2005; 71 FR 70464, Dec. 5, 2006]



Sec. 13.203  Separation of functions.

    (a) Civil penalty proceedings, including hearings, shall be 
prosecuted by an agency attorney.
    (b) An agency employee engaged in the performance of investigative 
or prosecutorial functions in a civil penalty action shall not, in that 
case or a factually-related case, participate or give advice in a 
decision by the administrative law judge or by the FAA decisionmaker on 
appeal, except as counsel or a witness in the public proceedings.
    (c) The Chief Counsel, the Deputy Chief Counsel for Policy and 
Adjudication, and the Assistant Chief Counsel for Litigation, or an 
attorney on the staff of the Assistant Chief Counsel for Litigation will 
advise the FAA decisionmaker regarding an initial decision or any appeal 
of a civil penalty action to the FAA decisionmaker.

[Amdt. 13-21, 55 FR 27575, July 3, 1990, as amended by Amdt. 13-24, 58 
FR 50241, Sept. 24, 1993; 70 FR 8238, Feb. 18, 2005]



Sec. 13.204  Appearances and rights of parties.

    (a) Any party may appear and be heard in person.
    (b) Any party may be accompanied, represented, or advised by an 
attorney

[[Page 59]]

or representative designated by the party and may be examined by that 
attorney or representative in any proceeding governed by this subpart. 
An attorney or representative who represents a party may file a notice 
of appearance in the action, in the manner provided in Sec. 13.210 of 
this subpart, and shall serve a copy of the notice of appearance on each 
party, in the manner provided in Sec. 13.211 of this subpart, before 
participating in any proceeding governed by this subpart. The attorney 
or representative shall include the name, address, and telephone number 
of the attorney or representative in the notice of appearance.
    (c) Any person may request a copy of a document upon payment of 
reasonable costs. A person may keep an original document, data, or 
evidence, with the consent of the administrative law judge, by 
substituting a legible copy of the document for the record.



Sec. 13.205  Administrative law judges.

    (a) Powers of an administrative law judge. In accordance with the 
rules of this subpart, an administrative law judge may:
    (1) Give notice of, and hold, prehearing conferences and hearings;
    (2) Administer oaths and affirmations;
    (3) Issue subpoenas authorized by law and issue notices of 
deposition requested by the parties;
    (4) Rule on offers of proof;
    (5) Receive relevant and material evidence;
    (6) Regulate the course of the hearing in accordance with the rules 
of this subpart;
    (7) Hold conferences to settle or to simplify the issues by consent 
of the parties;
    (8) Dispose of procedural motions and requests; and
    (9) Make findings of fact and conclusions of law, and issue an 
initial decision.
    (b) Limitations on the power of the administrative law judge. The 
administrative law judge shall not issue an order of contempt, award 
costs to any party, or impose any sanction not specified in this 
subpart. If the administrative law judge imposes any sanction not 
specified in this subpart, a party may file an interlocutory appeal of 
right with the FAA decisionmaker pursuant to Sec. 13.219(c)(4) of this 
subpart. This section does not preclude an administrative law judge from 
issuing an order that bars a person from a specific proceeding based on 
a finding of obstreperous or disruptive behavior in that specific 
proceeding.
    (c) Disqualification. The administrative law judge may disqualify 
himself or herself at any time. A party may file a motion, pursuant to 
Sec. 13.218(f)(6), requesting that an administrative law judge be 
disqualified from the proceedings.

[Amdt. 13-21, 55 FR 27575, July 3, 1990; 55 FR 29293, July 18, 1990]



Sec. 13.206  Intervention.

    (a) A person may submit a motion for leave to intervene as a party 
in a civil penalty action. Except for good cause shown, a motion for 
leave to intervene shall be submitted not later than 10 days before the 
hearing.
    (b) If the administrative law judge finds that intervention will not 
unduly broaden the issues or delay the proceedings, the administrative 
law judge may grant a motion for leave to intervene if the person will 
be bound by any order or decision entered in the action or the person 
has a property, financial, or other legitimate interest that may not be 
addressed adequately by the parties. The administrative law judge may 
determine the extent to which an intervenor may participate in the 
proceedings.



Sec. 13.207  Certification of documents.

    (a) Signature required. The attorney of record, the party, or the 
party's representative shall sign each document tendered for filing with 
the hearing docket clerk, the administrative law judge, the FAA 
decisionmaker on appeal, or served on each party.
    (b) Effect of signing a document. By signing a document, the 
attorney of record, the party, or the party's representative certifies 
that the attorney, the party, or the party's representative has read the 
document and, based on reasonable inquiry and to the best of that 
person's knowledge, information, and belief, the document is--
    (1) Consistent with these rules;

[[Page 60]]

    (2) Warranted by existing law or that a good faith argument exists 
for extension, modification, or reversal of existing law; and
    (3) Not unreasonable or unduly burdensome or expensive, not made to 
harass any person, not made to cause unnecessary delay, not made to 
cause needless increase in the cost of the proceedings, or for any other 
improper purpose.
    (c) Sanctions. If the attorney of record, the party, or the party's 
representative signs a document in violation of this section, the 
administrative law judge or the FAA decisionmaker shall:
    (1) Strike the pleading signed in violation of this section;
    (2) Strike the request for discovery or the discovery response 
signed in violation of this section and preclude further discovery by 
the party;
    (3) Deny the motion or request signed in violation of this section;
    (4) Exclude the document signed in violation of this section from 
the record;
    (5) Dismiss the interlocutory appeal and preclude further appeal on 
that issue by the party who filed the appeal until an initial decision 
has been entered on the record; or
    (6) Dismiss the appeal of the administrative law judge's initial 
decision to the FAA decisionmaker.



Sec. 13.208  Complaint.

    (a) Filing. The agency attorney shall file the original and one copy 
of the complaint with the hearing docket clerk, or may file a written 
motion pursuant to Sec. l3.218(f)(2)(i) of this subpart instead of 
filing a complaint, not later than 20 days after receipt by the agency 
attorney of a request for hearing. The agency attorney should suggest a 
location for the hearing when filing the complaint.
    (b) Service. An agency attorney shall personally deliver or mail a 
copy of the complaint on the respondent, the president of the 
corporation or company named as a respondent, or a person designated by 
the respondent to accept service of documents in the civil penalty 
action.
    (c) Contents. A complaint shall set forth the facts alleged, any 
regulation allegedly violated by the respondent, and the proposed civil 
penalty in sufficient detail to provide notice of any factual or legal 
allegation and proposed civil penalty.
    (d) Motion to dismiss allegations or complaint. Instead of filing an 
answer to the complaint, a respondent may move to dismiss the complaint, 
or that part of the complaint, alleging a violation that occurred on or 
after August 2, 1990, and more than 2 years before an agency attorney 
issued a notice of proposed civil penalty to the respondent.
    (1) An administrative law judge may not grant the motion and dismiss 
the complaint or part of the complaint if the administrative law judge 
finds that the agency has shown good cause for any delay in issuing the 
notice of proposed civil penalty.
    (2) If the agency fails to show good cause for any delay, an 
administrative law judge may dismiss the complaint, or that part of the 
complaint, alleging a violation that occurred more than 2 years before 
an agency attorney issued the notice of proposed civil penalty to the 
respondent.
    (3) A party may appeal the administrative law judge's ruling on the 
motion to dismiss the complaint or any part of the complaint in 
accordance with Sec. 13.219(b) of this subpart.

[Admt. 13-21, 55 FR 27575, July 3, 1990, as amended by Admt. 13-22, 55 
FR 31176, Aug. 1, 1990]



Sec. 13.209  Answer.

    (a) Writing required. A respondent shall file a written answer to 
the complaint, or may file a written motion pursuant to Sec. 13.208(d) 
or Sec. 13.218(f)(1-4) of this subpart instead of filing an answer, not 
later than 30 days after service of the complaint. The answer may be in 
the form of a letter but must be dated and signed by the person 
responding to the complaint. An answer may be typewritten or may be 
legibly handwritten.
    (b) Filing and address. A person filing an answer shall personally 
deliver or mail the original and one copy of the answer for filing with 
the hearing docket clerk, not later than 30 days after service of the 
complaint to the

[[Page 61]]

Hearing Docket at the appropriate address set forth in Sec. 13.210(a) 
of this subpart. The person filing an answer should suggest a location 
for the hearing when filing the answer.
    (c) Service. A person filing an answer shall serve a copy of the 
answer on the agency attorney who filed the complaint.
    (d) Contents. An answer shall specifically state any affirmative 
defense that the respondent intends to assert at the hearing. A person 
filing an answer may include a brief statement of any relief requested 
in the answer.
    (e) Specific denial of allegations required. A person filing an 
answer shall admit, deny, or state that the person is without sufficient 
knowledge or information to admit or deny, each numbered paragraph of 
the complaint. Any statement or allegation contained in the complaint 
that is not specifically denied in the answer may be deemed an admission 
of the truth of that allegation. A general denial of the complaint is 
deemed a failure to file an answer.
    (f) Failure to file answer. A person's failure to file an answer 
without good cause shall be deemed an admission of the truth of each 
allegation contained in the complaint.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979,as amended at 70 FR 8238, 
Feb. 18, 2005]



Sec. 13.210  Filing of documents.

    (a) Address and method of filing. A person tendering a document for 
filing shall personally deliver or mail the signed original and one copy 
of each document to the Hearing Docket using the appropriate address:
    (1) If delivery is in person, or via expedited courier service: 
Federal Aviation Administration, 600 Independence Avenue, SW., Wilbur 
Wright Building--Suite 2W1000, Washington, DC 20591; Attention: Hearing 
Docket Clerk, AGC-430.
    (2) If delivery is via U.S. Mail: Federal Aviation Administration, 
800 Independence Avenue, SW., Washington, DC 20591; Attention: Hearing 
Docket Clerk, AGC-430, Wilbur Wright Building--Suite 2W1000.
    (b) Date of filing. A document shall be considered to be filed on 
the date of personal delivery; or if mailed, the mailing date shown on 
the certificate of service, the date shown on the postmark if there is 
no certificate of service, or other mailing date shown by other evidence 
if there is no certificate of service or postmark.
    (c) Form. Each document shall be typewritten or legibly handwritten.
    (d) Contents. Unless otherwise specified in this subpart, each 
document must contain a short, plain statement of the facts on which the 
person's case rests and a brief statement of the action requested in the 
document.
    (e) Internet accessibility of documents filed in the Hearing Docket. 
(1) Unless protected from public disclosure by an order of the ALJ under 
13.226, all documents filed in the Hearing Docket are accessible through 
the Federal Docket Management System (FDMS): http://www.regulations.gov. 
To access a particular case file, use the FDMS number assigned to the 
case.
    (2) Decisions and orders issued by the Administrator in civil 
penalty cases, indexes of decisions, contact information for the FAA 
Hearing Docket and the administrative law judges, the rules of practice, 
and other information are available on the FAA civil penalty 
adjudication Web site at: http://www.faa.gov/about/office--org/
headquarters--offices/agc/pol--adjudication/AGC400/Civil--Penalty.

[Amdt. 13-21, 55 FR 27575, July 3, 1990; 55 FR 29293, July 18, 1990, as 
amended at 70 FR 8238, Feb. 18, 2005; 71 FR 70464, Dec. 5, 2006; 72 FR 
14668, Mar. 29, 2007; 72 FR 68474, Dec. 5, 2007]



Sec. 13.211  Service of documents.

    (a) General. A person shall serve a copy of any document filed with 
the Hearing Docket on each party at the time of filing. Service on a 
party's attorney of record or a party's designated representative may be 
considered adequate service on the party.
    (b) Type of service. A person may serve documents by personal 
delivery or by mail.
    (c) Certificate of service. A person may attach a certificate of 
service to a document tendered for filing with the hearing docket clerk. 
A certificate of service shall consist of a statement, dated and signed 
by the person filing the document, that the document was personally 
delivered or mailed to each party on a specific date.

[[Page 62]]

    (d) Date of service. The date of service shall be the date of 
personal delivery; or if mailed, the mailing date shown on the 
certificate of service, the date shown on the postmark if there is no 
certificate of service, or other mailing date shown by other evidence if 
there is no certificate of service or postmark.
    (e) Additional time after service by mail. Whenever a party has a 
right or a duty to act or to make any response within a prescribed 
period after service by mail, or on a date certain after service by 
mail, 5 days shall be added to the prescribed period.
    (f) Service by the administrative law judge. The administrative law 
judge shall serve a copy of each document including, but not limited to, 
notices of prehearing conferences and hearings, rulings on motions, 
decisions, and orders, upon each party to the proceedings by personal 
delivery or by mail.
    (g) Valid service. A document that was properly addressed, was sent 
in accordance with this subpart, and that was returned, that was not 
claimed, or that was refused, is deemed to have been served in 
accordance with this subpart. The service shall be considered valid as 
of the date and the time that the document was deposited with a contract 
or express messenger, the document was mailed, or personal delivery of 
the document was refused.
    (h) Presumption of service. There shall be a presumption of service 
where a party or a person, who customarily receives mail, or receives it 
in the ordinary course of business, at either the person's residence or 
the person's principal place of business, acknowledges receipt of the 
document.



Sec. 13.212  Computation of time.

    (a) This section applies to any period of time prescribed or allowed 
by this subpart, by notice or order of the administrative law judge, or 
by any applicable statute.
    (b) The date of an act, event, or default, after which a designated 
time period begins to run, is not included in a computation of time 
under this subpart.
    (c) The last day of a time period is included in a computation of 
time unless it is a Saturday, Sunday, or a legal holiday. If the last 
day of the time period is a Saturday, Sunday, or legal holiday, the time 
period runs until the end of the next day that is not a Saturday, 
Sunday, or legal holiday.



Sec. 13.213  Extension of time.

    (a) Oral requests. The parties may agree to extend for a reasonable 
period the time for filing a document under this subpart. If the parties 
agree, the administrative law judge shall grant one extension of time to 
each party. The party seeking the extension of time shall submit a draft 
order to the administrative law judge to be signed by the administrative 
law judge and filed with the hearing docket clerk. The administrative 
law judge may grant additional oral requests for an extension of time 
where the parties agree to the extension.
    (b) Written motion. A party shall file a written motion for an 
extension of time with the administrative law judge not later than 7 
days before the document is due unless good cause for the late filing is 
shown. A party filing a written motion for an extension of time shall 
serve a copy of the motion on each party. The administrative law judge 
may grant the extension of time if good cause for the extension is 
shown.
    (c) Failure to rule. If the administrative law judge fails to rule 
on a written motion for an extension of time by the date the document 
was due, the motion for an extension of time is deemed granted for no 
more than 20 days after the original date the document was to be filed.



Sec. 13.214  Amendment of pleadings.

    (a) Filing and service. A party shall file the amendment with the 
administrative law judge and shall serve a copy of the amendment on all 
parties to the proceeding.
    (b) Time. A party shall file an amendment to a complaint or an 
answer within the following:
    (1) Not later than 15 days before the scheduled date of a hearing, a 
party may amend a complaint or an answer without the consent of the 
administrative law judge.

[[Page 63]]

    (2) Less than 15 days before the scheduled date of a hearing, the 
administrative law judge may allow amendment of a complaint or an answer 
only for good cause shown in a motion to amend.
    (c) Responses. The administrative law judge shall allow a reasonable 
time, but not more than 20 days from the date of filing, for other 
parties to respond if an amendment to a complaint, answer, or other 
pleading has been filed with the administrative law judge.



Sec. 13.215  Withdrawal of complaint or request for hearing.

    At any time before or during a hearing, an agency attorney may 
withdraw a complaint or a party may withdraw a request for a hearing 
without the consent of the administrative law judge. If an agency 
attorney withdraws the complaint or a party withdraws the request for a 
hearing and the answer, the administrative law judge shall dismiss the 
proceedings under this subpart with prejudice.



Sec. 13.216  Waivers.

    Waivers of any rights provided by statute or regulation shall be in 
writing or by stipulation made at a hearing and entered into the record. 
The parties shall set forth the precise terms of the waiver and any 
conditions.



Sec. 13.217  Joint procedural or discovery schedule.

    (a) General. The parties may agree to submit a schedule for filing 
all prehearing motions, a schedule for conducting discovery in the 
proceedings, or a schedule that will govern all prehearing motions and 
discovery in the proceedings.
    (b) Form and content of schedule. If the parties agree to a joint 
procedural or discovery schedule, one of the parties shall file the 
joint schedule with the administrative law judge, setting forth the 
dates to which the parties have agreed, and shall serve a copy of the 
joint schedule on each party.
    (1) The joint schedule may include, but need not be limited to, 
requests for discovery, any objections to discovery requests, responses 
to discovery requests to which there are no objections, submission of 
prehearing motions, responses to prehearing motions, exchange of 
exhibits to be introduced at the hearing, and a list of witnesses that 
may be called at the hearing.
    (2) Each party shall sign the original joint schedule to be filed 
with the administrative law judge.
    (c) Time. The parties may agree to submit all prehearing motions and 
responses and may agree to close discovery in the proceedings under the 
joint schedule within a reasonable time before the date of the hearing, 
but not later than 15 days before the hearing.
    (d) Order establishing joint schedule. The administrative law judge 
shall approve the joint schedule filed by the parties. One party shall 
submit a draft order establishing a joint schedule to the administrative 
law judge to be signed by the administrative law judge and filed with 
the hearing docket clerk.
    (e) Disputes. The administrative law judge shall resolve disputes 
regarding discovery or disputes regarding compliance with the joint 
schedule as soon as possible so that the parties may continue to comply 
with the joint schedule.
    (f) Sanctions for failure to comply with joint schedule. If a party 
fails to comply with the administrative law judge's order establishing a 
joint schedule, the administrative law judge may direct that party to 
comply with a motion to discovery request or, limited to the extent of 
the party's failure to comply with a motion or discovery request, the 
administrative law judge may:
    (1) Strike that portion of a party's pleadings;
    (2) Preclude prehearing or discovery motions by that party;
    (3) Preclude admission of that portion of a party's evidence at the 
hearing, or
    (4) Preclude that portion of the testimony of that party's witnesses 
at the hearing.



Sec. 13.218  Motions.

    (a) General. A party applying for an order or ruling not 
specifically provided in this subpart shall do so by motion. A party 
shall comply with the requirements of this section when filing a motion 
with the administrative law

[[Page 64]]

judge. A party shall serve a copy of each motion on each party.
    (b) Form and contents. A party shall state the relief sought by the 
motion and the particular grounds supporting that relief. If a party has 
evidence in support of a motion, the party shall attach any supporting 
evidence, including affidavits, to the motion.
    (c) Filing of motions. A motion made prior to the hearing must be in 
writing. Unless otherwise agreed by the parties or for good cause shown, 
a party shall file any prehearing motion, and shall serve a copy on each 
party, not later than 30 days before the hearing. Motions introduced 
during a hearing may be made orally on the record unless the 
administrative law judge directs otherwise.
    (d) Answers to motions. Any party may file an answer, with 
affidavits or other evidence in support of the answer, not later than 10 
days after service of a written motion on that party. When a motion is 
made during a hearing, the answer may be made at the hearing on the 
record, orally or in writing, within a reasonable time determined by the 
administrative law judge.
    (e) Rulings on motions. The administrative law judge shall rule on 
all motions as follows:
    (1) Discovery motions. The administrative law judge shall resolve 
all pending discovery motions not later than 10 days before the hearing.
    (2) Prehearing motions. The administrative law judge shall resolve 
all pending prehearing motions not later than 7 days before the hearing. 
If the administrative law judge issues a ruling or order orally, the 
administrative law judge shall serve a written copy of the ruling or 
order, within 3 days, on each party. In all other cases, the 
administrative law judge shall issue rulings and orders in writing and 
shall serve a copy of the ruling or order on each party.
    (3) Motions made during the hearing. The administrative law judge 
may issue rulings and orders on motions made during the hearing orally. 
Oral rulings or orders on motions must be made on the record.
    (f) Specific motions. A party may file the following motions with 
the administrative law judge:
    (1) Motion to dismiss for insufficiency. A respondent may file a 
motion to dismiss the complaint for insufficiency instead of filing an 
answer. If the administrative law judge denies the motion to dismiss the 
complaint for insufficiency, the respondent shall file an answer not 
later than 10 days after service of the administrative law judge's 
denial of the motion.A motion to dismiss the complaint for insufficiency 
must show that the complaint fails to state a violation of a provision 
of the Federal aviation statute listed in the first sentence in 49 
U.S.C. 46301(d)(2) or in 49 U.S.C. 47531, or any implementing rule, 
regulation, or order, or a violation of the Federal hazardous materials 
transportation statute, 49 U.S.C. 5121-5128, or any implementing rule, 
regulation or order.
    (2) Motion to dismiss. A party may file a motion to dismiss, 
specifying the grounds for dismissal. If an administrative law judge 
grants a motion to dismiss in part, a party may appeal the 
administrative law judge's ruling on the motion to dismiss under Sec. 
13.219(b) of this subpart.
    (i) Motion to dismiss a request for a hearing. An agency attorney 
may file a motion to dismiss a request for a hearing instead of filing a 
complaint. If the motion to dismiss is not granted, the agency attorney 
shall file the complaint and shall serve a copy of the complaint on each 
party not later than 10 days after service of the administrative law 
judge's ruling or order on the motion to dismiss. If the motion to 
dismiss is granted and the proceedings are terminated without a hearing, 
the respondent may file an appeal pursuant to Sec. 13.233 of this 
subpart. If required by the decision on appeal, the agency attorney 
shall file a complaint and shall serve a copy of the complaint on each 
party not later than 10 days after service of the decision on appeal.
    (ii) Motion to dismiss a complaint. A respondent may file a motion 
to dismiss a complaint instead of filing an answer. If the motion to 
dismiss is not granted, the respondent shall file an answer and shall 
serve a copy of the answer on each party not later than 10 days after 
service of the administrative law judge's ruling or order on the motion 
to dismiss. If the motion to dismiss is

[[Page 65]]

granted and the proceedings are terminated without a hearing, the agency 
attorney may file an appeal pursuant to Sec. 13.233 of this subpart. If 
required by the decision on appeal, the respondent shall file an answer 
and shall serve a copy of the answer on each party not later than 10 
days after service of the decision on appeal.
    (3) Motion for more definite statement. A party may file a motion 
for more definite statement of any pleading which requires a response 
under this subpart. A party shall set forth, in detail, the indefinite 
or uncertain allegations contained in a complaint or response to any 
pleading and shall submit the details that the party believes would make 
the allegation or response definite and certain.
    (i) Complaint. A respondent may file a motion requesting a more 
definite statement of the allegations contained in the complaint instead 
of filing an answer. If the administrative law judge grants the motion, 
the agency attorney shall supply a more definite statement not later 
than 15 days after service of the ruling granting the motion. If the 
agency attorney fails to supply a more definite statement, the 
administrative law judge shall strike the allegations in the complaint 
to which the motion is directed. If the administrative law judge denies 
the motion, the respondent shall file an answer and shall serve a copy 
of the answer on each party not later than 10 days after service of the 
order of denial.
    (ii) Answer. An agency attorney may file a motion requesting a more 
definite statement if an answer fails to respond clearly to the 
allegations in the complaint. If the administrative law judge grants the 
motion, the respondent shall supply a more definite statement not later 
than 15 days after service of the ruling on the motion. If the 
respondent fails to supply a more definite statement, the administrative 
law judge shall strike those statements in the answer to which the 
motion is directed. The respondent's failure to supply a more definite 
statement may be deemed an admission of unanswered allegations in the 
complaint.
    (4) Motion to strike. Any party may make a motion to strike any 
insufficient allegation or defense, or any redundant, immaterial, or 
irrelevant matter in a pleading. A party shall file a motion to strike 
with the administrative law judge and shall serve a copy on each party 
before a response is required under this subpart or, if a response is 
not required, not later than 10 days after service of the pleading.
    (5) Motion for decision. A party may make a motion for decision, 
regarding all or any part of the proceedings, at any time before the 
administrative law judge has issued an initial decision in the 
proceedings. The administrative law judge shall grant a party's motion 
for decision if the pleadings, depositions, answers to interrogatories, 
admissions, matters that the administrative law judge has officially 
noticed, or evidence introduced during the hearing show that there is no 
genuine issue of material fact and that the party making the motion is 
entitled to a decision as a matter of law. The party making the motion 
for decision has the burden of showing that there is no genuine issue of 
material fact disputed by the parties.
    (6) Motion for disqualification. A party may file a motion for 
disqualification with the administrative law judge and shall serve a 
copy on each party. A party may file the motion at any time after the 
administrative law judge has been assigned to the proceedings but shall 
make the motion before the administrative law judge files an initial 
decision in the proceedings.
    (i) Motion and supporting affidavit. A party shall state the grounds 
for disqualification, including, but not limited to, personal bias, 
pecuniary interest, or other factors showing disqualification, in the 
motion for disqualification. A party shall submit an affidavit with the 
motion for disqualification that sets forth, in detail, the matters 
alleged to constitute grounds for disqualification.
    (ii) Answer. A party shall respond to the motion for 
disqualification not later than 5 days after service of the motion for 
disqualification.
    (iii) Decision on motion for disqualification. The administrative 
law judge shall render a decision on the motion for disqualification not 
later than 15 days after the motion has been filed. If the 
administrative law judge finds that

[[Page 66]]

the motion for disqualification and supporting affidavit show a basis 
for disqualification, the administrative law judge shall withdraw from 
the proceedings immediately. If the administrative law judge finds that 
disqualification is not warranted, the administrative law judge shall 
deny the motion and state the grounds for the denial on the record. If 
the administrative law judge fails to rule on a party's motion for 
disqualification within 15 days after the motion has been filed, the 
motion is deemed granted.
    (iv) Appeal. A party may appeal the administrative law judge's 
denial of the motion for disqualification in accordance with Sec. 
13.219(b) of this subpart.

[Amdt. 13-21, 55 FR 27575, July 3, 1990, as amended by 71 FR 70464, Dec. 
5, 2006]



Sec. 13.219  Interlocutory appeals.

    (a) General. Unless otherwise provided in this subpart, a party may 
not appeal a ruling or decision of the administrative law judge to the 
FAA decisionmaker until the initial decision has been entered on the 
record. A decision or order of the FAA decisionmaker on the 
interlocutory appeal does not constitute a final order of the 
Administrator for the purposes of judicial appellate review as provided 
in Sec. 13.235 of this subpart.
    (b) Interlocutory appeal for cause. If a party files a written 
request for an interlocutory appeal for cause with the administrative 
law judge, or orally requests an interlocutory appeal for cause, the 
proceedings are stayed until the administrative law judge issues a 
decision on the request. If the administrative law judge grants the 
request, the proceedings are stayed until the FAA decisionmaker issues a 
decision on the interlocutory appeal. The administrative law judge shall 
grant an interlocutory appeal for cause if a party shows that delay of 
the appeal would be detrimental to the public interest or would result 
in undue prejudice to any party.
    (c) Interlocutory appeals of right. If a party notifies the 
administrative law judge of an interlocutory appeal of right, the 
proceedings are stayed until the FAA decisionmaker issues a decision on 
the interlocutory appeal. A party may file an interlocutory appeal with 
the FAA decisionmaker, without the consent of the administrative law 
judge, before an initial decision has been entered in the case of:
    (1) A ruling or order by the administrative law judge barring a 
person from the proceedings.
    (2) Failure of the administrative law judge to dismiss the 
proceedings in accordance with Sec. 13.215 of this subpart.
    (3) A ruling or order by the administrative law judge in violation 
of Sec. 13.205(b) of this subpart.
    (d) Procedure. A party shall file a notice of interlocutory appeal, 
with supporting documents, with the FAA decisionmaker and the hearing 
docket clerk, and shall serve a copy of the notice and supporting 
documents on each party and the administrative law judge, not later than 
10 days after the administrative law judge's decision forming the basis 
of an interlocutory appeal of right or not later than 10 days after the 
administrative law judge's decision granting an interlocutory appeal for 
cause, whichever is appropriate. A party shall file a reply brief, if 
any, with the FAA decisionmaker and serve a copy of the reply brief on 
each party, not later than 10 days after service of the appeal brief. 
The FAA decisionmaker shall render a decision on the interlocutory 
appeal, on the record and as a part of the decision in the proceedings, 
within a reasonable time after receipt of the interlocutory appeal.
    (e) The FAA decisionmaker may reject frivolous, repetitive, or 
dilatory appeals, and may issue an order precluding one or more parties 
from making further interlocutory appeals in a proceeding in which there 
have been frivolous, repetitive, or dilatory interlocutory appeals.

[Amdt. 13-21, 55 FR 27575, July 3, 1990, as amended by Amdt. 13-23, 55 
FR 45983, Oct. 31, 1990; 71 FR 70464, Dec. 5, 2006]



Sec. 13.220  Discovery.

    (a) Initiation of discovery. Any party may initiate discovery 
described in this section, without the consent or approval of the 
administrative law judge, at any time after a complaint has been filed 
in the proceedings.

[[Page 67]]

    (b) Methods of discovery. The following methods of discovery are 
permitted under this section: depositions on oral examination or written 
questions of any person; written interrogatories directed to a party; 
requests for production of documents or tangible items to any person; 
and requests for admission by a party. A party is not required to file 
written interrogatories and responses, requests for production of 
documents or tangible items and responses, and requests for admission 
and response with the administrative law judge or the hearing docket 
clerk. In the event of a discovery dispute, a party shall attach a copy 
of these documents in support of a motion made under this section.
    (c) Service on the agency. A party shall serve each discovery 
request directed to the agency or any agency employee on the agency 
attorney of record.
    (d) Time for response to discovery requests. Unless otherwise 
directed by this subpart or agreed by the parties, a party shall respond 
to a request for discovery, including filing objections to a request for 
discovery, not later than 30 days of service of the request.
    (e) Scope of discovery. Subject to the limits on discovery set forth 
in paragraph (f) of this section, a party may discover any matter that 
is not privileged and that is relevant to the subject matter of the 
proceeding. A party may discover information that relates to the claim 
or defense of any party including the existence, description, nature, 
custody, condition, and location of any document or other tangible item 
and the identity and location of any person having knowledge of 
discoverable matter. A party may discover facts known, or opinions held, 
by an expert who any other party expects to call to testify at the 
hearing. A party has no ground to object to a discovery request on the 
basis that the information sought would not be admissible at the hearing 
if the information sought during discovery is reasonably calculated to 
lead to the discovery of admissible evidence.
    (f) Limiting discovery. The administrative law judge shall limit the 
frequency and extent of discovery permitted by this section if a party 
shows that--
    (1) The information requested is cumulative or repetitious;
    (2) The information requested can be obtained from another less 
burdensome and more convenient source;
    (3) The party requesting the information has had ample opportunity 
to obtain the information through other discovery methods permitted 
under this section; or
    (4) The method or scope of discovery requested by the party is 
unduly burdensome or expensive.
    (g) Confidential orders. A party or person who has received a 
discovery request for information that is related to a trade secret, 
confidential or sensitive material, competitive or commercial 
information, proprietary data, or information on research and 
development, may file a motion for a confidential order with the 
administrative law judge and shall serve a copy of the motion for a 
confidential order on each party.
    (1) The party or person making the motion must show that the 
confidential order is necessary to protect the information from 
disclosure to the public.
    (2) If the administrative law judge determines that the requested 
material is not necessary to decide the case, the administrative law 
judge shall preclude any inquiry into the matter by any party.
    (3) If the administrative law judge determines that the requested 
material may be disclosed during discovery, the administrative law judge 
may order that the material may be discovered and disclosed under 
limited conditions or may be used only under certain terms and 
conditions.
    (4) If the administrative law judge determines that the requested 
material is necessary to decide the case and that a confidential order 
is warranted, the administrative law judge shall provide:
    (i) An opportunity for review of the document by the parties off the 
record;
    (ii) Procedures for excluding the information from the record; and
    (iii) Order that the parties shall not disclose the information in 
any manner and the parties shall not use the information in any other 
proceeding.

[[Page 68]]

    (h) Protective orders. A party or a person who has received a 
request for discovery may file a motion for protective order with the 
administrative law judge and shall serve a copy of the motion for 
protective order on each party. The party or person making the motion 
must show that the protective order is necessary to protect the party or 
the person from annoyance, embarrassment, oppression, or undue burden or 
expense. As part of the protective order, the administrative law judge 
may:
    (1) Deny the discovery request;
    (2) Order that discovery be conducted only on specified terms and 
conditions, including a designation of the time or place for discovery 
or a determination of the method of discovery; or
    (3) Limit the scope of discovery or preclude any inquiry into 
certain matters during discovery.
    (i) Duty to supplement or amend responses. A party who has responded 
to a discovery request has a duty to supplement or amend the response, 
as soon as the information is known, as follows:
    (1) A party shall supplement or amend any response to a question 
requesting the identity and location of any person having knowledge of 
discoverable matters.
    (2) A party shall supplement or amend any response to a question 
requesting the identity of each person who will be called to testify at 
the hearing as an expert witness and the subject matter and substance of 
that witness' testimony.
    (3) A party shall supplement or amend any response that was 
incorrect when made or any response that was correct when made but is no 
longer correct, accurate, or complete.
    (j) Depositions. The following rules apply to depositions taken 
pursuant to this section:
    (1) Form. A deposition shall be taken on the record and reduced to 
writing. The person being deposed shall sign the deposition unless the 
parties agree to waive the requirement of a signature.
    (2) Administration of oaths. Within the United States, or a 
territory or possession subject to the jurisdiction of the United 
States, a party shall take a deposition before a person authorized to 
administer oaths by the laws of the United States or authorized by the 
law of the place where the examination is held. In foreign countries, a 
party shall take a deposition in any manner allowed by the Federal Rules 
of Civil Procedure.
    (3) Notice of deposition. A party shall serve a notice of 
deposition, stating the time and place of the deposition and the name 
and address of each person to be examined, on the person to be deposed, 
on the administrative law judge, on the hearing docket clerk, and on 
each party not later than 7 days before the deposition. A party may 
serve a notice of deposition less than 7 days before the deposition only 
with consent of the administrative law judge. If a subpoena duces tecum 
is to be served on the person to be examined, the party shall attach a 
copy of the subpoena duces tecum that describes the materials to be 
produced at the deposition to the notice of deposition.
    (4) Use of depositions. A party may use any part or all of a 
deposition at a hearing authorized under this subpart only upon a 
showing of good cause. The deposition may be used against any party who 
was present or represented at the deposition or who had reasonable 
notice of the deposition.
    (k) Interrogatories. A party, the party's attorney, or the party's 
representative may sign the party's responses to interrogatories. A 
party shall answer each interrogatory separately and completely in 
writing. If a party objects to an interrogatory, the party shall state 
the objection and the reasons for the objection. An opposing party may 
use any part or all of a party's responses to interrogatories at a 
hearing authorized under this subpart to the extent that the response is 
relevant, material, and not repetitious.
    (1) A party shall not serve more than 30 interrogatories to each 
other party. Each subpart of an interrogatory shall be counted as a 
separate interrogatory.
    (2) A party shall file a motion for leave to serve additional 
interrogatories on a party with the administrative law judge before 
serving additional interrogatories on a party. The administrative law 
judge shall grant the motion only if the party shows good cause for the 
party's failure to inquire about the information previously and that

[[Page 69]]

the information cannot reasonably be obtained using less burdensome 
discovery methods or be obtained from other sources.
    (l) Requests for admission. A party may serve a written request for 
admission of the truth of any matter within the scope of discovery under 
this section or the authenticity of any document described in the 
request. A party shall set forth each request for admission separately. 
A party shall serve copies of documents referenced in the request for 
admission unless the documents have been provided or are reasonably 
available for inspection and copying.
    (1) Time. A party's failure to respond to a request for admission, 
in writing and signed by the attorney or the party, not later than 30 
days after service of the request, is deemed an admission of the truth 
of the statement or statements contained in the request for admission. 
The administrative law judge may determine that a failure to respond to 
a request for admission is not deemed an admission of the truth if a 
party shows that the failure was due to circumstances beyond the control 
of the party or the party's attorney.
    (2) Response. A party may object to a request for admission and 
shall state the reasons for objection. A party may specifically deny the 
truth of the matter or describe the reasons why the party is unable to 
truthfully deny or admit the matter. If a party is unable to deny or 
admit the truth of the matter, the party shall show that the party has 
made reasonable inquiry into the matter or that the information known 
to, or readily obtainable by, the party is insufficient to enable the 
party to admit or deny the matter. A party may admit or deny any part of 
the request for admission. If the administrative law judge determines 
that a response does not comply with the requirements of this rule or 
that the response is insufficient, the matter is deemed admitted.
    (3) Effect of admission. Any matter admitted or deemed admitted 
under this section is conclusively established for the purpose of the 
hearing and appeal.
    (m) Motion to compel discovery. A party may make a motion to compel 
discovery if a person refuses to answer a question during a deposition, 
a party fails or refuses to answer an interrogatory, if a person gives 
an evasive or incomplete answer during a deposition or when responding 
to an interrogatory, or a party fails or refuses to produce documents or 
tangible items. During a deposition, the proponent of a question may 
complete the deposition or may adjourn the examination before making a 
motion to compel if a person refuses to answer.
    (n) Failure to comply with a discovery order or order to compel. If 
a party fails to comply with a discovery order or an order to compel, 
the administrative law judge, limited to the extent of the party's 
failure to comply with the discovery order or motion to compel, may:
    (1) Strike that portion of a party's pleadings;
    (2) Preclude prehearing or discovery motions by that party;
    (3) Preclude admission of that portion of a party's evidence at the 
hearing; or
    (4) Preclude that portion of the testimony of that party's witnesses 
at the hearing.

[Amdt. 13-21, 55 FR 27575, July 3, 1990, as amended by Amdt. 13-23, 55 
FR 45983, Oct. 31, 1990]



Sec. 13.221  Notice of hearing.

    (a) Notice. The administrative law judge shall give each party at 
least 60 days notice of the date, time, and location of the hearing.
    (b) Date, time, and location of the hearing. The administrative law 
judge to whom the proceedings have been assigned shall set a reasonable 
date, time, and location for the hearing. The administrative law judge 
shall consider the need for discovery and any joint procedural or 
discovery schedule submitted by the parties when determining the hearing 
date. The administrative law judge shall give due regard to the 
convenience of the parties, the location where the majority of the 
witnesses reside or work, and whether the location is served by a 
scheduled air carrier.
    (c) Earlier hearing. With the consent of the administrative law 
judge, the parties may agree to hold the hearing

[[Page 70]]

on an earlier date than the date specified in the notice of hearing.



Sec. 13.222  Evidence.

    (a) General. A party is entitled to present the party's case or 
defense by oral, documentary, or demonstrative evidence, to submit 
rebuttal evidence, and to conduct any cross-examination that may be 
required for a full and true disclosure of the facts.
    (b) Admissibility. A party may introduce any oral, documentary, or 
demonstrative evidence in support of the party's case or defense. The 
administrative law judge shall admit any oral, documentary, or 
demonstrative evidence introduced by a party but shall exclude 
irrelevant, immaterial, or unduly repetitious evidence.
    (c) Hearsay evidence. Hearsay evidence is admissible in proceedings 
governed by this subpart. The fact that evidence submitted by a party is 
hearsay goes only to the weight of the evidence and does not affect its 
admissibility.



Sec. 13.223  Standard of proof.

    The administrative law judge shall issue an initial decision or 
shall rule in a party's favor only if the decision or ruling is 
supported by, and in accordance with, the reliable, probative, and 
substantial evidence contained in the record. In order to prevail, the 
party with the burden of proof shall prove the party's case or defense 
by a preponderance of reliable, probative, and substantial evidence.



Sec. 13.224  Burden of proof.

    (a) Except in the case of an affirmative defense, the burden of 
proof is on the agency.
    (b) Except as otherwise provided by statute or rule, the proponent 
of a motion, request, or order has the burden of proof.
    (c) A party who has asserted an affirmative defense has the burden 
of proving the affirmative defense.



Sec. 13.225  Offer of proof.

    A party whose evidence has been excluded by a ruling of the 
administrative law judge may offer the evidence for the record on 
appeal.



Sec. 13.226  Public disclosure of evidence.

    (a) The administrative law judge may order that any information 
contained in the record be withheld from public disclosure. Any person 
may object to disclosure of information in the record by filing a 
written motion to withhold specific information with the administrative 
law judge and serving a copy of the motion on each party. The party 
shall state the specific grounds for nondisclosure in the motion.
    (b) The administrative law judge shall grant the motion to withhold 
information in the record if, based on the motion and any response to 
the motion, the administrative law judge determines that disclosure 
would be detrimental to aviation safety, disclosure would not be in the 
public interest, or that the information is not otherwise required to be 
made available to the public.



Sec. 13.227  Expert or opinion witnesses.

    An employee of the agency may not be called as an expert or opinion 
witness, for any party other than the FAA, in any proceeding governed by 
this subpart. An employee of a respondent may not be called by an agency 
attorney as an expert or opinion witness for the FAA in any proceeding 
governed by this subpart to which the respondent is a party.



Sec. 13.228  Subpoenas.

    (a) Request for subpoena. A party may obtain a subpoena to compel 
the attendance of a witness at a deposition or hearing or to require the 
production of documents or tangible items from the hearing docket clerk. 
The hearing docket clerk shall deliver the subpoena, signed by the 
hearing docket clerk or an administrative law judge but otherwise in 
blank, to the party. The party shall complete the subpoena, stating the 
title of the action and the date and time for the witness' attendance or 
production of documents or items. The party who obtained the subpoena 
shall serve the subpoena on the witness.
    (b) Motion to quash or modify the subpoena. A party, or any person 
upon whom a subpoena has been served, may file a motion to quash or 
modify the subpoena with the administrative law

[[Page 71]]

judge at or before the time specified in the subpoena for compliance. 
The applicant shall describe, in detail, the basis for the application 
to quash or modify the supoena including, but not limited to, a 
statement that the testimony, document, or tangible evidence is not 
relevant to the proceeding, that the subpoena is not reasonably tailored 
to the scope of the proceeding, or that the subpoena is unreasonable and 
oppressive. A motion to quash or modify the subpoena will stay the 
effect of the subpoena pending a decision by the administrative law 
judge on the motion.
    (c) Enforcement of subpoena. Upon a showing that a person has failed 
or refused to comply with a subpoena, a party may apply to the local 
federal district court to seek judicial enforcement of the subpoena in 
accordance with 49 U.S.C. 46104 in cases under the Federal aviation 
statute.

[Amdt. 13-21, 55 FR 27575, July 3, 1990, as amended at 71 FR 70465, Dec. 
5, 2006]



Sec. 13.229  Witness fees.

    (a) General. Unless otherwise authorized by the administrative law 
judge, the party who applies for a subpoena to compel the attendance of 
a witness at a deposition or hearing, or the party at whose request a 
witness appears at a deposition or hearing, shall pay the witness fees 
described in this section.
    (b) Amount. Except for an employee of the agency who appears at the 
direction of the agency, a witness who appears at a deposition or 
hearing is entitled to the same fees and mileage expenses as are paid to 
a witness in a court of the United States in comparable circumstances.



Sec. 13.230  Record.

    (a) Exclusive record. The transcript of all testimony in the 
hearing, all exhibits received into evidence, and all motions, 
applications, requests, and rulings shall constitute the exclusive 
record for decision of the proceedings and the basis for the issuance of 
any orders in the proceeding. Any proceedings regarding the 
disqualification of an administrative law judge shall be included in the 
record.
    (b) Examination and copying of record. Any person may examine the 
record at the Hearing Docket, Federal Aviation Administration, 600 
Independence Avenue, SW., Wilbur Wright Building--Room 2014, Washington, 
DC 20591. Documents may also be examined and copied at the U.S. 
Department of Transportation, Docket Operations, West Building Ground 
Floor, Room W12-140, 1200 New Jersey Avenue, SE., Washington, DC 20590. 
Any person may have a copy of the record after payment of reasonable 
costs to copy the record.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended at 70 FR 8238, 
Feb. 18, 2005; 72 FR 68474, Dec. 5, 2007]



Sec. 13.231  Argument before the administrative law judge.

    (a) Arguments during the hearing. During the hearing, the 
administrative law judge shall give the parties a reasonable opportunity 
to present arguments on the record supporting or opposing motions, 
objections, and rulings if the parties request an opportunity for 
argument. The administrative law judge may request written arguments 
during the hearing if the administrative law judge finds that submission 
of written arguments would be reasonable.
    (b) Final oral argument. At the conclusion of the hearing and before 
the administrative law judge issues an initial decision in the 
proceedings, the parties are entitled to submit oral proposed findings 
of fact and conclusions of law, exceptions to rulings of the 
administrative law judge, and supporting arguments for the findings, 
conclusions, or exceptions. At the conclusion of the hearing, a party 
may waive final oral argument.
    (c) Posthearing briefs. The administrative law judge may request 
written posthearing briefs before the administrative law judge issues an 
initial decision in the proceedings if the administrative law judge 
finds that submission of written arguments would be reasonable. If a 
party files a written posthearing brief, the party shall include 
proposed findings of fact and conclusions of law, exceptions to rulings 
of the administrative law judge, and supporting arguments for the 
findings, conclusions, or exceptions. The administrative law judge shall 
give the parties a reasonable opportunity, not more than 30 days after 
receipt of the

[[Page 72]]

transcript, to prepare and submit the briefs.



Sec. 13.232  Initial decision.

    (a) Contents. The administrative law judge shall issue an initial 
decision at the conclusion of the hearing. In each oral or written 
decision, the administrative law judge shall include findings of fact 
and conclusions of law, and the grounds supporting those findings and 
conclusions, upon all material issues of fact, the credibility of 
witnesses, the applicable law, any exercise of the administrative law 
judge's discretion, the amount of any civil penalty found appropriate by 
the administrative law judge, and a discussion of the basis for any 
order issued in the proceedings. The administrative law judge is not 
required to provide a written explanation for rulings on objections, 
procedural motions, and other matters not directly relevant to the 
substance of the initial decision. If the administrative law judge 
refers to any previous unreported or unpublished initial decision, the 
administrative law judge shall make copies of that initial decision 
available to all parties and the FAA decisionmaker.
    (b) Oral decision. Except as provided in paragraph (c) of this 
section, at the conclusion of the hearing, the administrative law judge 
shall issue the initial decision and order orally on the record.
    (c) Written decision. The administrative law judge may issue a 
written initial decision not later than 30 days after the conclusion of 
the hearing or submission of the last posthearing brief if the 
administrative law judge finds that issuing a written initial decision 
is reasonable. The administrative law judge shall serve a copy of any 
written initial decision on each party.
    (d) Order assessing civil penalty. Unless appealed pursuant to Sec. 
13.233 of this subpart, the initial decision issued by the 
administrative law judge shall be considered an order assessing civil 
penalty if the administrative law judge finds that an alleged violation 
occurred and determines that a civil penalty, in an amount found 
appropriate by the administrative law judge, is warranted.



Sec. 13.233  Appeal from initial decision.

    (a) Notice of appeal. A party may appeal the initial decision, and 
any decision not previously appealed pursuant to Sec. 13.219, by filing 
a notice of appeal with the FAA decisionmaker. A party must file the 
notice of appeal in the FAA Hearing Docket using the appropriate address 
listed in Sec. 13.210(a). A party shall file the notice of appeal not 
later than 10 days after entry of the oral initial decision on the 
record or service of the written initial decision on the parties and 
shall serve a copy of the notice of appeal on each party.
    (b) Issues on appeal. In any appeal from a decision of an 
administrative law judge, the FAA decisionmaker considers only the 
following issues:
    (1) Whether each finding of fact is supported by a preponderance of 
reliable, probative, and substantial evidence;
    (2) Whether each conclusion of law is made in accordance with 
applicable law, precedent, and public policy; and
    (3) Whether the administrative law judge committed any prejudicial 
errors that support the appeal.
    (c) Perfecting an appeal. Unless otherwise agreed by the parties, a 
party shall perfect an appeal, not later than 50 days after entry of the 
oral initial decision on the record or service of the written initial 
decision on the party, by filing an appeal brief with the FAA 
decisionmaker.
    (1) Extension of time by agreement of the parties. The parties may 
agree to extend the time for perfecting the appeal with the consent of 
the FAA decisionmaker. If the FAA decisionmaker grants an extension of 
time to perfect the appeal, the appellate docket clerk shall serve a 
letter confirming the extension of time on each party.
    (2) Written motion for extension. If the parties do not agree to an 
extension of time for perfecting an appeal, a party desiring an 
extension of time may file a written motion for an extension with the 
FAA decisionmaker and shall serve a copy of the motion on each party. 
The FAA decisionmaker may grant an extension if good cause for the 
extension is shown in the motion.

[[Page 73]]

    (d) Appeal briefs. A party shall file the appeal brief with the FAA 
decisionmaker and shall serve a copy of the appeal brief on each party.
    (1) A party shall set forth, in detail, the party's specific 
objections to the initial decision or rulings in the appeal brief. A 
party also shall set forth, in detail, the basis for the appeal, the 
reasons supporting the appeal, and the relief requested in the appeal. 
If the party relies on evidence contained in the record for the appeal, 
the party shall specifically refer to the pertinent evidence contained 
in the transcript in the appeal brief.
    (2) The FAA decisionmaker may dismiss an appeal, on the FAA 
decisionmaker's own initiative or upon motion of any other party, where 
a party has filed a notice of appeal but fails to perfect the appeal by 
timely filing an appeal brief with the FAA decisionmaker.
    (e) Reply brief. Unless otherwise agreed by the parties, any party 
may file a reply brief with the FAA decisionmaker not later than 35 days 
after the appeal brief has been served on that party. The party filing 
the reply brief shall serve a copy of the reply brief on each party. If 
the party relies on evidence contained in the record for the reply, the 
party shall specifically refer to the pertinent evidence contained in 
the transcript in the reply brief.
    (1) Extension of time by agreement of the parties. The parties may 
agree to extend the time for filing a reply brief with the consent of 
the FAA decisionmaker. If the FAA decisionmaker grants an extension of 
time to file the reply brief, the appellate docket clerk shall serve a 
letter confirming the extension of time on each party.
    (2) Written motion for extension. If the parties do not agree to an 
extension of time for filing a reply brief, a party desiring an 
extension of time may file a written motion for an extension with the 
FAA decisionmaker and shall serve a copy of the motion on each party. 
The FAA decisionmaker may grant an extension if good cause for the 
extension is shown in the motion.
    (f) Other briefs. The FAA decisionmaker may allow any person to 
submit an amicus curiae brief in an appeal of an initial decision. A 
party may not file more than one appeal brief or reply brief. A party 
may petition the FAA decisionmaker, in writing, for leave to file an 
additional brief and shall serve a copy of the petition on each party. 
The party may not file the additional brief with the petition. The FAA 
decisionmaker may grant leave to file an additional brief if the party 
demonstrates good cause for allowing additional argument on the appeal. 
The FAA decisionmaker will allow a reasonable time for the party to file 
the additional brief.
    (g) Number of copies. A party shall file the original appeal brief 
or the original reply brief, and two copies of the brief, with the FAA 
decisionmaker.
    (h) Oral argument. The FAA decisionmaker has sole discretion to 
permit oral argument on the appeal. On the FAA decisionmaker's own 
initiative or upon written motion by any party, the FAA decisionmaker 
may find that oral argument will contribute substantially to the 
development of the issues on appeal and may grant the parties an 
opportunity for oral argument.
    (i) Waiver of objections on appeal. If a party fails to object to 
any alleged error regarding the proceedings in an appeal or a reply 
brief, the party waives any objection to the alleged error. The FAA 
decisionmaker is not required to consider any objection in an appeal 
brief or any argument in the reply brief if a party's objection is based 
on evidence contained on the record and the party does not specifically 
refer to the pertinent evidence from the record in the brief.
    (j) FAA decisionmaker's decision on appeal. The FAA decisionmaker 
will review the record, the briefs on appeal, and the oral argument, if 
any, when considering the issues on appeal. The FAA decisionmaker may 
affirm, modify, or reverse the initial decision, make any necessary 
findings, or may remand the case for any proceedings that the FAA 
decisionmaker determines may be necessary.
    (1) The FAA decisionmaker may raise any issue, on the FAA 
decisionmaker's own initiative, that is required for proper disposition 
of the proceedings. The FAA decisionmaker will give the parties a 
reasonable opportunity to submit arguments on the new issues

[[Page 74]]

before making a decision on appeal. If an issue raised by the FAA 
decisionmaker requires the consideration of additional testimony or 
evidence, the FAA decisionmaker will remand the case to the 
administrative law judge for further proceedings and an initial decision 
related to that issue. If an issue raised by the FAA decisionmaker is 
solely an issue of law or the issue was addressed at the hearing but was 
not raised by a party in the briefs on appeal, a remand of the case to 
the administrative law judge for further proceedings is not required but 
may be provided in the discretion of the FAA decisionmaker.
    (2) The FAA decisionmaker will issue the final decision and order of 
the Administrator on appeal in writing and will serve a copy of the 
decision and order on each party. Unless a petition for review is filed 
pursuant to Sec. 13.235, a final decision and order of the 
Administrator shall be considered an order assessing civil penalty if 
the FAA decisionmaker finds that an alleged violation occurred and a 
civil penalty is warranted.
    (3) A final decision and order of the Administrator after appeal is 
precedent in any other civil penalty action. Any issue, finding or 
conclusion, order, ruling, or initial decision of an administrative law 
judge that has not been appealed to the FAA decisionmaker is not 
precedent in any other civil penalty action.

[Amdt. 13-21, 55 FR 27575, July 3, 1990, as amended by Amdt. 13-32; 69 
FR 59498, Oct. 4, 2004; 70 FR 13345, Mar. 21, 2005]



Sec. 13.234  Petition to reconsider or modify a final decision and
order of the FAA decisionmaker on appeal.

    (a) General. Any party may petition the FAA decisionmaker to 
reconsider or modify a final decision and order issued by the FAA 
decisionmaker on appeal from an initial decision. A party shall file a 
petition to reconsider or modify with the FAA decisionmaker not later 
than 30 days after service of the FAA decisionmaker's final decision and 
order on appeal and shall serve a copy of the petition on each party. 
The FAA decisionmaker will not reconsider or modify an initial decision 
and order issued by an administrative law judge that has not been 
appealed by any party to the FAA decisionmaker.
    (b) Form and number of copies. A party shall file a petition to 
reconsider or modify, in writing, with the FAA decisionmaker. The party 
shall file the original petition with the FAA decisionmaker and shall 
serve a copy of the petition on each party.
    (c) Contents. A party shall state briefly and specifically the 
alleged errors in the final decision and order on appeal, the relief 
sought by the party, and the grounds that support, the petition to 
reconsider or modify.
    (1) If the petition is based, in whole or in part, on allegations 
regarding the consequences of the FAA decisionmaker's decision, the 
party shall describe these allegations and shall describe, and support, 
the basis for the allegations.
    (2) If the petition is based, in whole or in part, on new material 
not previously raised in the proceedings, the party shall set forth the 
new material and include affidavits of prospective witnesses and 
authenticated documents that would be introduced in support of the new 
material. The party shall explain, in detail, why the new material was 
not discovered through due diligence prior to the hearing.
    (d) Repetitious and frivolous petitions. The FAA decisionmaker will 
not consider repetitious or frivolous petitions. The FAA decisionmaker 
may summarily dismiss repetitious or frivolous petitions to reconsider 
or modify.
    (e) Reply petitions. Any other party may reply to a petition to 
reconsider or modify, not later than 10 days after service of the 
petition on that party, by filing a reply with the FAA decisionmaker. A 
party shall serve a copy of the reply on each party.
    (f) Effect of filing petition. Unless otherwise ordered by the FAA 
decisionmaker, filing of a petition pursuant to this section will not 
stay or delay the effective date of the FAA decisionmaker's final 
decision and order on appeal and shall not toll the time allowed for 
judicial review.
    (g) FAA decisionmaker's decision on petition. The FAA decisionmaker 
has sole discretion to grant or deny a petition to reconsider or modify. 
The FAA decisionmaker will grant or deny a petition

[[Page 75]]

to reconsider or modify within a reasonable time after receipt of the 
petition or receipt of the reply petition, if any. The FAA decisionmaker 
may affirm, modify, or reverse the final decision and order on appeal, 
or may remand the case for any proceedings that the FAA decisionmaker 
determines may be necessary.

[Amdt. 13-21, 55 FR 27575, July 3, 1990; 55 FR 29293, July 18, 1990; 
Amdt. 13-23, 55 FR 45983, Oct. 31, 1990]



Sec. 13.235  Judicial review of a final decision and order.

    (a) In cases under the Federal aviation statute, a party may seek 
judicial review of a final decision and order of the Administrator, as 
provided in 49 U.S.C. 46110(a), and, as applicable, in 49 U.S.C. 
46301(d)(7)(D)(iii), 46301(g), or 47532.
    (b) In cases under the Federal hazardous materials transportation 
statute, a party may seek judicial review of a final decision and order 
of the Administrator, as provided in 49 U.S.C. 5127.
    (c) A party seeking judicial review of a final order issued by the 
Administrator may file a petition for review in the United States Court 
of Appeals for the District of Columbia Circuit or in the United States 
Court of Appeals for the circuit in which the party resides or has its 
principal place of business.
    (d) The party must file the petition for review no later than 60 
days after service of the Administrator's final decision and order.

[Doc. No. FAA-2006-26477, 71 FR 70465, Dec. 5, 2006]



          Subpart H_Civil Monetary Penalty Inflation Adjustment

    Source: Docket No. 28762, 61 FR 67445, Dec. 20, 1996, unless 
otherwise noted.



Sec. 13.301  Scope and purpose.

    (a) This subpart provides a mechanism for the regular adjustment for 
inflation of civil monetary penalties in conformity with the Federal 
Civil Penalties Inflation Adjustment Act of 1990, 28 U.S.C. 2461 (note), 
as amended by the Debt Collection Improvement Act of 1996, Public Law 
104-134, April 26, 1996, in order to maintain the deterrent effect of 
civil monetary penalties and to promote compliance with the law. This 
subpart also sets out the current adjusted maximum civil monetary 
penalties or range of minimum and maximum civil monetary penalties for 
each statutory civil penalty subject to the FAA's jurisdiction.
    (b) Each adjustment to the maximum civil monetary penalty or the 
range of minimum and maximum civil monetary penalties, as applicable, 
made in accordance with this subpart applies prospectively from the date 
it becomes effective to actions initiated under this part, 
notwithstanding references to a specific maximum civil monetary penalty 
or range of minimum and maximum civil monetary penalties contained 
elsewhere in this part.



Sec. 13.303  Definitions.

    (a) Civil Monetary Penalty means any penalty, fine, or other 
sanction that:
    (1) Is for a specific monetary amount as provided by Federal law or 
has a maximum amount provided by Federal law;
    (2) Is assessed or enforced by the FAA pursuant to Federal law; and
    (3) Is assessed or enforced pursuant to an administrative proceeding 
or a civil action in the Federal courts.
    (b) Consumer Price Index means the Consumer Price Index for all 
urban consumers published by the Department of Labor.



Sec. 13.305  Cost of living adjustments of civil monetary penalties.

    (a) Except for the limitation to the initial adjustment to statutory 
maximum civil monetary penalties or range of minimum and maximum civil 
monetary penalties set forth in paragraph (c) of this section, the 
inflation adjustment under this subpart is determined by increasing the 
maximum civil monetary penalty or range of minimum and maximum civil 
monetary penalty for each civil monetary penalty by the cost-of-living 
adjustment. Any increase determined under paragraph (a) of this section 
is rounded to the nearest:
    (1) Multiple of $10 in the case of penalties less than or equal to 
$100;

[[Page 76]]

    (2) Multiple of $100 in the case of penalties greater than $100 but 
less than or equal to $1,000;
    (3) Multiple of $1,000 in the case of penalties greater than $1,000 
but less than or equal to $10,000;
    (4) Multiple of $5,000 in the case of penalties greater than $10,000 
but less than or equal to $100,000;
    (5) Multiple of $10,000 in the case of penalties greater than 
$100,000 but less than or equal to $200,000; and
    (6) Multiple of $25,000 in the case of penalties greater than 
$200,000.
    (b) For purposes of paragraph (a) of this section, the term ``cost-
of-living adjustment'' means the percentage (if any) for each civil 
monetary penalty by which the Consumer Price Index for the month of June 
of the calendar year preceding the adjustment exceeds the Consumer Price 
Index for the month of June of the calendar year in which the amount of 
such civil monetary penalty was last set or adjusted pursuant to law.
    (c) Limitation on initial adjustment. The initial adjustment of a 
civil monetary penalty under this subpart does not exceed 10 percent of 
the civil penalty amount.
    (d) Inflation adjustment. Minimum and maximum civil monetary 
penalties within the jurisdiction of the FAA are adjusted for inflation 
as follows: Minimum and Maximum Civil Penalties-Adjusted for Inflation.

               Table of Minimum and Maximum Civil Monetary Penalty Amounts for Certain Violations Occurring On or After December 29, 2010
--------------------------------------------------------------------------------------------------------------------------------------------------------
                                                                                                               Maximum penalty
                                      Civil monetary penalty     Minimum penalty        New or adjusted      amount when last set     New or adjusted
      United States Code cite              description                amount            minimum penalty      or adjusted pursuant     maximum penalty
                                                                                             amount                 to law                 amount
--------------------------------------------------------------------------------------------------------------------------------------------------------
49 U.S.C. 5123(a), subparagraph (1)  Violation of hazardous   $250 per violation,    No change............  $50,000 per            $55,000 per
                                      materials                reset 8/10/2005.                              violation, reset 8/    violation.
                                      transportation law.                                                    10/2005.
49 U.S.C. 5123(a), subparagraph (2)  Violation of hazardous   $250 per violation,    No change............  $100,000 per           $110,000 per
                                      materials                reset 8/10/2005.                              violation, set 8/10/   violation.
                                      transportation law                                                     2005.
                                      resulting in death,
                                      serious illness,
                                      severe injury, or
                                      substantial property
                                      destruction.
49 U.S.C. 5123(a), subparagraph (3)  Violation of hazardous   $450 per violation,    No change............  $50,000 per            $55,000 per
                                      materials                set 8/10/2005.                                violation, set 8/10/   violation.
                                      transportation law                                                     2005.
                                      relating to training.
49 U.S.C. 46301(a)(1)..............  Violation by a person    N/A..................  N/A..................  $25,000 per            $27,500 per
                                      other than an                                                          violation, reset 12/   violation.
                                      individual or small                                                    12/2003.
                                      business concern under
                                      49 CFR 46301(a)(1)(A)
                                      or (B).
49 U.S.C. 46301(a)(1)..............  Violation by an airman   N/A..................  N/A..................  $1,100 per violation,  No change.
                                      serving as an airman                                                   reset 12/12/2003.
                                      under 49 U.S.C.
                                      46301(a)(1)(A) or (B)
                                      (but not covered by
                                      46301(a)(5)(A) or (B).
49 U.S.C. 46301(a)(1)..............  Violation by an          N/A..................  N/A..................  $1,100 per violation,  No change.
                                      individual or small                                                    reset 12/12/2003.
                                      business concern under
                                      49 U.S.C.
                                      46301(a)(1)(A) or (B)
                                      (but not covered in 49
                                      U.S.C. 46301(a)(5)).
49 U.S.C. 46301(a)(3)..............  Violation of 49 U.S.C.   N/A..................  N/A..................  Increase above         No change.
                                      47107(b) (or any                                                       otherwise applicable
                                      assurance made under                                                   maximum amount not
                                      such section) or 49                                                    to exceed 3 times
                                      U.S.C. 47133.                                                          the amount of
                                                                                                             revenues that are
                                                                                                             used in violation of
                                                                                                             such section.
49 U.S.C. 46301(a)(5)(A)...........  Violation by an          N/A..................  N/A..................  $11,000 per            No change.
                                      individual or small                                                    violation, adjusted
                                      business concern                                                       6/15/2006.
                                      (except an airman
                                      serving as an airman)
                                      under 49 U.S.C.
                                      46301(a)(5)(A)(i) or
                                      (ii).
49 U.S.C. 46301(a)(5)(B)(i)........  Violation by an          N/A..................  N/A..................  $11,000 per            No change.
                                      individual or small                                                    violation, adjusted
                                      business concern                                                       6/15/2006.
                                      related to the
                                      transportation of
                                      hazardous materials.

[[Page 77]]

 
49 U.S.C. 46301(a)(5)(B)(ii).......  Violation by an          N/A..................  N/A..................  $11,000 per            No change.
                                      individual or small                                                    violation, adjusted
                                      business concern                                                       6/16/2006.
                                      related to the
                                      registration or
                                      recordation under 49
                                      U.S.C. chapter 441, of
                                      an aircraft not used
                                      to provide air
                                      transportation.
49 U.S.C. 46301(a)(5)(B)(iii)......  Violation by an          N/A..................  N/A..................  $11,000 per            No change.
                                      individual or small                                                    violation, adjusted
                                      business concern of 49                                                 6/15/2006.
                                      U.S.C. 44718(d),
                                      relating to limitation
                                      on construction or
                                      establishment of
                                      landfills.
49 U.S.C. 46301(a)(5)(B)(iv).......  Violation by an          N/A..................  N/A..................  $11,000 per            No change.
                                      individual or small                                                    violation, adjusted
                                      business concern of 49                                                 6/15/2006.
                                      U.S.C. 44725, relating
                                      to the safe disposal
                                      of life-limited
                                      aircraft parts.
49 U.S.C. 46301(b).................  Tampering with a smoke   N/A..................  N/A..................  $2,200 per violation,  $3,200 per violation.
                                      alarm device.                                                          adjusted 1/21/1997.
49 U.S.C. 46302....................  Knowingly providing      N/A..................  N/A..................  $11,000 per            $16,000 per
                                      false information                                                      violation, adjusted    violation.
                                      about alleged                                                          1/21/1997.
                                      violation involving
                                      the special aircraft
                                      jurisdiction of the
                                      United States.
49 U.S.C. 46318....................  Interference with cabin  N/A..................  N/A..................  $27,500, adjusted 6/   No change.
                                      or flight crew.                                                        15/2006.
49 U.S.C. 46319....................  Permanent closure of an  N/A..................  N/A..................  $11,000 per day,       No change.
                                      airport without                                                        adjusted 6/15/2006.
                                      providing sufficient
                                      notice.
49 U.S.C. 47531....................  Violation of 49 U.S.C.   N/A..................  N/A..................  See 49 U.S.C.          No change.
                                      47528-47530, relating                                                  46301(a)(1)(A) and
                                      to the prohibition of                                                  (a)(5), above.
                                      operating certain
                                      aircraft not complying
                                      with stage 3 noise
                                      levels.
--------------------------------------------------------------------------------------------------------------------------------------------------------


[61 FR 67445, Dec. 20, 1996, as amended by Amdt. 13-28, 62 FR 4134, Jan. 
29, 1997; 67 FR 6366, Feb. 11, 2002; Amdt. 13-33, 71 FR 28522, May 16, 
2006; 71 FR 47077, Aug. 16, 2006; 71 FR 52407, Sept. 6, 2006; 75 FR 
72938, Nov. 29, 2010]



         Subpart I_Flight Operational Quality Assurance Programs



Sec. 13.401  Flight Operational Quality Assurance Program: Prohibition
against use of data for enforcement purposes.

    (a) Applicability. This section applies to any operator of an 
aircraft who operates such aircraft under an approved Flight Operational 
Quality Assurance (FOQA) program.
    (b) Definitions. For the purpose of this section, the terms--
    (1) Flight Operational Quality Assurance (FOQA) program means an 
FAA-approved program for the routine collection and analysis of digital 
flight data gathered during aircraft operations, including data 
currently collected pursuant to existing regulatory provisions, when 
such data is included in an approved FOQA program.
    (2) FOQA data means any digital flight data that has been collected 
from an individual aircraft pursuant to an FAA-approved FOQA program, 
regardless of the electronic format of that data.
    (3) Aggregate FOQA data means the summary statistical indices that 
are associated with FOQA event categories, based on an analysis of FOQA 
data from multiple aircraft operations.
    (c) Requirements. In order for paragraph (e) of this section to 
apply, the operator must submit, maintain, and adhere to a FOQA 
Implementation and Operation Plan that is approved by the

[[Page 78]]

Administrator and which contains the following elements:
    (1) A description of the operator's plan for collecting and 
analyzing flight recorded data from line operations on a routine basis, 
including identification of the data to be collected;
    (2) Procedures for taking corrective action that analysis of the 
data indicates is necessary in the interest of safety;
    (3) Procedures for providing the FAA with aggregate FOQA data;
    (4) Procedures for informing the FAA as to any corrective action 
being undertaken pursuant to paragraph (c)(2) of this section.
    (d) Submission of aggregate data. The operator will provide the FAA 
with aggregate FOQA data in a form and manner acceptable to the 
Administrator.
    (e) Enforcement. Except for criminal or deliberate acts, the 
Administrator will not use an operator's FOQA data or aggregate FOQA 
data in an enforcement action against that operator or its employees 
when such FOQA data or aggregate FOQA data is obtained from a FOQA 
program that is approved by the Administrator.
    (f) Disclosure. FOQA data and aggregate FOQA data, if submitted in 
accordance with an order designating the information as protected under 
part 193 of this chapter, will be afforded the nondisclosure protections 
of part 193 of this chapter.
    (g) Withdrawal of program approval. The Administrator may withdraw 
approval of a previously approved FOQA program for failure to comply 
with the requirements of this chapter. Grounds for withdrawal of 
approval may include, but are not limited to--
    (1) Failure to implement corrective action that analysis of 
available FOQA data indicates is necessary in the interest of safety; or
    (2) Failure to correct a continuing pattern of violations following 
notice by the agency; or also
    (3) Willful misconduct or willful violation of the FAA regulations 
in this chapter.

[Doc. No. FAA-2000-7554, 66 FR 55048, Oct. 31, 2001; Amdt. 13-30, 67 FR 
31401, May 9, 2002]



PART 14_RULES IMPLEMENTING THE EQUAL ACCESS TO JUSTICE ACT OF 1980--
Table of Contents



                      Subpart A_General Provisions

Sec.
14.01 Purpose of these rules.
14.02 Proceedings covered.
14.03 Eligibility of applicants.
14.04 Standards for awards.
14.05 Allowance fees and expenses.

             Subpart B_Information Required From Applicants

14.10 Contents of application.
14.11 Net worth exhibit.
14.12 Documentation of fees and expenses.

            Subpart C_Procedures for Considering Applications

14.20 When an application may be filed.
14.21 Filing and service of documents.
14.22 Answer to application.
14.23 Reply.
14.24 Comments by other parties.
14.25 Settlement.
14.26 Further proceedings.
14.27 Decision.
14.28 Review by FAA decisionmaker.
14.29 Judicial review.
14.30 Payment of award.

    Authority: 5 U.S.C. 504; 49 U.S.C. 106(f), 40113, 46104 and 47122.

    Source: Docket No. 25958, 54 FR 46199, Nov. 1, 1989, unless 
otherwise noted.



                      Subpart A_General Provisions



Sec. 14.01  Purpose of these rules.

    The Equal Access to Justice Act, 5 U.S.C. 504 (the Act), provides 
for the award of attorney fees and other expenses to eligible 
individuals and entities who are parties to certain administrative 
proceedings (adversary adjudications) before the Federal Aviation 
Administration (FAA). An eligible party may receive an award when it 
prevails over the FAA, unless the agency's position in the proceeding 
was substantially justified or special circumstances make an award 
unjust. The rules in this part describe the parties eligible for awards 
and the proceedings that are covered. They also explain how to apply for 
awards, and the procedures and standards that the FAA Decisionmaker will 
use to make them. As

[[Page 79]]

used hereinafter, the term ``agency'' applies to the FAA.



Sec. 14.02  Proceedings covered.

    (a) The Act applies to certain adversary adjudications conducted by 
the FAA under 49 CFR part 17 and the Acquisition Management System 
(AMS). These are adjudications under 5 U.S.C. 554, in which the position 
of the FAA is represented by an attorney or other representative who 
enters an appearance and participates in the proceeding. This subpart 
applies to proceedings under 49 U.S.C. 46301, 46302, and 46303 and to 
the Default Adjudicative Process under part 17 of this chapter and the 
AMS.
    (b) If a proceeding includes both matters covered by the Act and 
matters specifically excluded from coverage, any award made will include 
only fees and expenses related to covered issues.
    (c) Fees and other expenses may not be awarded to a party for any 
portion of the adversary adjudication in which such party has 
unreasonably protracted the proceedings.

[54 FR 46199, Nov. 1, 1989, as amended by Amdt. 14-03, 64 FR 32935, June 
18, 1999]



Sec. 14.03  Eligibility of applicants.

    (a) To be eligible for an award of attorney fees and other expenses 
under the Act, the applicant must be a party to the adversary 
adjudication for which it seeks an award. The term ``party'' is defined 
in 5 U.S.C. 504(b)(1)(B) and 5 U.S.C. 551(3). The applicant must show 
that it meets all conditions or eligibility set out in this subpart.
    (b) The types of eligible applicants are as follows:
    (1) An individual with a net worth of not more than $2 million at 
the time the adversary adjudication was initiated;
    (2) The sole owner of an unincorporated business who has a net worth 
of not more than $7 million, including both personal and business 
interests, and not more than 500 employees at the time the adversary 
adjudication was initiated;
    (3) A charitable or other tax-exempt organization described in 
section 501(c)(3) of the Internal Revenue Code (26 U.S.C. 501(c)(3)) 
with not more than 500 employees at the time the adversary adjudication 
was initiated; and
    (4) A cooperative association as defined in section 15(a) of the 
Agricultural Marketing Act (12 U.S.C. 1141j(a)) with not more than 500 
employees at the time the adversary adjudication was initiated; and
    (5) Any other partnership, corporation, association, or public or 
private organization with a net worth of not more than $7 million and 
not more than 500 employees at the time the adversary adjudication was 
initiated.
    (c) For the purpose of eligibility, the net worth and number of 
employees of an applicant shall be determined as of the date the 
proceeding was initiated.
    (d) An applicant who owns an unincorporated business will be 
considered an ``individual'' rather than a ``sole owner of an 
unincorporated business'' if the issues on which the applicant prevails 
are related primarily to personal interests rather than to business 
interest.
    (e) The employees of an applicant include all persons who regularly 
perform services for remuneration for the applicant, under the 
applicant's direction and control. Part-time employees shall be included 
on a proportional basis.
    (f) The net worth and number of employees of the applicant and all 
of its affiliates shall be aggregated to determine eligibility. Any 
individual, corporation, or other entity that directly or indirectly 
controls or owns a majority of the voting shares or other interest of 
the applicant, or any corporation or other entity of which the applicant 
directly or indirectly owns or controls a majority of the voting shares 
or other interest, will be considered an affiliate for purposes of this 
part, unless the ALJ or adjudicative officer determines that such 
treatment would be unjust and contrary to the purposes of the Act in 
light of the actual relationship between the affiliated entities. In 
addition, the ALJ or adjudicative officer may determine that financial 
relationships of the applicant, other than those described in this 
paragraph, constitute special circumstances that would make an award 
unjust.
    (g) An applicant that participates in a proceeding primarily on 
behalf of one

[[Page 80]]

or more other persons or entities that would be ineligible if not itself 
eligible for an award.

[54 FR 46199, Nov. 1, 1989, as amended by Amdt. 14-03, 64 FR 32935, June 
18, 1999]



Sec. 14.04  Standards for awards.

    (a) A prevailing applicant may receive an award for attorney fees 
and other expenses incurred in connection with a proceeding, or in a 
significant and discrete substantive portion of the proceeding, unless 
the position of the agency over which the applicant has prevailed was 
substantially justified. Whether or not the position of the FAA was 
substantially justified shall be determined on the basis of the record 
(including the record with respect to the action or failure to act by 
the agency upon which the civil action is based) which was made in the 
civil action for which fees and other expenses are sought. The burden of 
proof that an award should not be made to an eligible prevailing 
applicant is on the agency counsel, who may avoid an award by showing 
that the agency's position was reasonable in law and fact.
    (b) An award will be reduced or denied if the applicant has unduly 
or unreasonably protracted the proceeding or if special circumstances 
make the award sought unjust.



Sec. 14.05  Allowance fees and expenses.

    (a) Awards will be based on rates customarily charged by persons 
engaged in the business of acting as attorneys, agents, and expert 
witnesses, even if the services were made available without charge or at 
a reduced rate to the applicant.
    (b) No award for the fee of an attorney or agent under this part may 
exceed $125 per hour, or such rate as prescribed by 5 U.S.C. 504. No 
award to compensate an expert witness may exceed the highest rate at 
which the agency pays expert witnesses. However, an award may also 
include the reasonable expenses of the attorney, agent, or witness as a 
separate item, if the attorney, agent, or witness ordinarily charges 
clients separately for such expenses.
    (c) In determining the reasonableness of the fee sought for an 
attorney, agent, or expert witness, the ALJ or adjudicative officer 
shall consider the following:
    (1) If the attorney, agent, or witness is in private practice, his 
or her customary fee for similar services, or if an employee of the 
applicant, the fully allocated cost of the services;
    (2) The prevailing rate for similar services in the community in 
which the attorney, agent, or witness ordinarily performs services;
    (3) The time actually spent in the representation of the applicant;
    (4) The time reasonably spent in light of the difficulty or 
complexity of the issues in the proceeding; and
    (5) Such other factors as may bear on the value of the services 
provided.
    (d) The reasonable cost of any study, analysis, engineering report, 
test, project, or similar matter prepared on behalf of a party may be 
awarded, to the extent that the charge for the service does not exceed 
the prevailing rate for similar services, and the study or other matter 
was necessary for preparation of the applicant's case.
    (e) Fees may be awarded only for work performed after the issuance 
of a complaint, or in the Default Adjudicative Process for a protest or 
contract dispute under part 17 of this chapter and the AMS.

[Amdt. 13-18, 53 FR 34655, Sept. 7, 1988, as amended by Amdt. 14-1, 55 
FR 15131, Apr. 20, 1990; Amdt. 14-03, 64 FR 32935, June 18, 1999]



             Subpart B_Information Required From Applicants



Sec. 14.10  Contents of application.

    (a) An application for an award of fees and expenses under the Act 
shall identify the applicant and the proceeding for which an award is 
sought. The application shall show that the applicant has prevailed and 
identify the position of the agency in the proceeding that the applicant 
alleges was not substantially justified. Unless the applicant is an 
individual, the application shall also state the number of employees of 
the applicant and describe briefly the type and purpose of its 
organization or business.
    (b) The application shall also include a statement that the 
applicant's net worth does not exceed $2 million (if an individual) or 
$7 million (for all other

[[Page 81]]

applicants, including their affiliates) at the time the adversary 
adjudication was initiated. However, an applicant may omit this 
statement if:
    (1) It attaches a copy of a ruling by the Internal Revenue Service 
that it qualifies as an organization described in section 501(c)(3) of 
the Internal Revenue Code (26 U.S.C. 501(c)(3)), or in the case of a 
tax-exempt organization not required to obtain a ruling from the 
Internal Revenue Service on its exempt status, a statement that 
describes the basis for the applicant's belief that it qualifies under 
such section; or
    (2) It states that it is a cooperative association as defined in 
section 15(a) of the Agricultural Marketing Act (12 U.S.C. 1141j(a)).
    (c) The application shall state the amount of fees and expenses for 
which an award is sought.
    (d) The application may also include any other matters that the 
applicant wishes this agency to consider in determining whether and in 
what amount an award should be made.
    (e) The application shall be signed by the applicant or an 
authorized officer or attorney for the applicant. It shall also contain 
or be accompanied by a written verification under oath or under penalty 
of perjury that the information provided in the application is true and 
correct.
    (f) If the applicant is a partnership, corporation, association, 
organization, or sole owner of an unincorporated business, the 
application shall state that the applicant did not have more than 500 
employees at the time the adversary adjudication was initiated, giving 
the number of its employees and describing briefly the type and purpose 
of its organization or business.



Sec. 14.11  Net worth exhibit.

    (a) Each applicant except a qualified tax-exempt organization or 
cooperative association must provide with its application a detailed 
exhibit showing the net worth of the applicant and any affiliates when 
the proceeding was initiated. If any individual, corporation, or other 
entity directly or indirectly controls or owns a majority of the voting 
shares or other interest of the applicant, or if the applicant directly 
or indirectly owns or controls a majority of the voting shares or other 
interest of any corporation or other entity, the exhibit must include a 
showing of the net worth of all such affiliates or of the applicant 
including the affiliates. The exhibit may be in any form convenient to 
the applicant that provides full disclosure of the applicant's and its 
affiliates' assets and liabilities and is sufficient to determine 
whether the applicant qualifies under the standards in this part. The 
administrative law judge may require an applicant to file additional 
information to determine the eligibility for an award.
    (b) The net worth exhibit shall describe any transfers of assets 
from, or obligations incurred by, the applicant or any affiliate, 
occurring in the one-year period prior to the date on which the 
proceeding was initiated, that reduced the net worth of the applicant 
and its affiliates below the applicable net worth ceiling. If there were 
no such transactions, the applicant shall so state.
    (c) Ordinarily, the net worth exhibit will be included in the public 
record of the proceeding. However, an applicant that objects to public 
disclosure of the net worth exhibit, or any part of it, may submit that 
portion of the exhibit directly to the ALJ or adjudicative officer in a 
sealed envelope labeled ``Confidential Financial Information,'' 
accompanied by a motion to withhold the information.
    (1) The motion shall describe the information sought to be withheld 
and explain, in detail, why it should be exempt under applicable law or 
regulation, why public disclosure would adversely affect the applicant, 
and why disclosure is not required in the public interest.
    (2) The net worth exhibit shall be served on the FAA counsel, but 
need not be served on any other party to the proceeding.
    (3) If the ALJ or adjudicative officer finds that the net worth 
exhibit, or any part of it, should not be withheld from disclosure, it 
shall be placed in the public record of the proceeding. Otherwise, any 
request to inspect or copy

[[Page 82]]

the exhibit shall be disposed of in accordance with the FAA's 
established procedures.

[54 FR 46199, Nov. 1, 1989, as amended by Amdt. 14-03, 64 FR 32935, June 
18, 1999]



Sec. 14.12  Documentation of fees and expenses.

    The application shall be accompanied by full documentation of the 
fees and expenses, including the cost of any study, analysis, 
engineering report, test, project or similar matter, for which an award 
is sought. A separate itemized statement shall be submitted for each 
professional firm or individual whose services are covered by the 
application, showing the hours spent in connection with the proceedings 
by each individual, a description of the specific services performed, 
the rate at which each fee has been computed, any expenses for which 
reimbursement is sought, the total amount claimed, and the total amount 
paid or payable by the applicant or by any other person or entity for 
the services provided. The administrative law judge may require the 
applicant to provide vouchers, receipts, or other substantiation for any 
expenses claimed.



            Subpart C_Procedures for Considering Applications



Sec. 14.20  When an application may be filed.

    (a) An application may be filed whenever the applicant has prevailed 
in the proceeding, but in no case later than 30 days after the FAA 
Decisionmaker's final disposition of the proceeding, or service of the 
order of the Administrator in a proceeding under the AMS.
    (b) If review or reconsideration is sought or taken of a decision to 
which an applicant believes it has prevailed, proceedings for the award 
of fees shall be stayed pending final disposition of the underlying 
controversy.
    (c) For purposes of this part, final disposition means the later of:
    (1) Under part 17 of this chapter and the AMS, the date on which the 
order of the Administrator is served;
    (2) The date on which an unappealed initial decision becomes 
administratively final;
    (3) Issuance of an order disposing of any petitions for 
reconsideration of the FAA Decisionmaker's final order in the 
proceeding;
    (4) If no petition for reconsideration is filed, the last date on 
which such a petition could have been filed; or
    (5) Issuance of a final order or any other final resolution of a 
proceeding, such as a settlement or voluntary dismissal, which is not 
subject to a petition for reconsideration.

[54 FR 46199, Nov. 1, 1989, as amended by Amdt. 14-03, 64 FR 32936, June 
18, 1999]



Sec. 14.21  Filing and service of documents.

    Any application for an award or other pleading or document related 
to an application shall be filed and served on all parties to the 
proceeding in the same manner as other pleadings in the proceeding, 
except as provided in Sec. 14.11(b) for confidential financial 
information. Where the proceeding was held under part 17 of this chapter 
and the AMS, the application shall be filed with the FAA's attorney and 
with the Office of Dispute Resolution for Acquisition.

[Doc. No. FAA-1998-4379, 64 FR 32936, June 18, 1999]



Sec. 14.22  Answer to application.

    (a) Within 30 days after service of an application, counsel 
representing the agency against which an award is sought may file an 
answer to the application. Unless agency counsel requests an extension 
of time for filing or files a statement of intent to negotiate under 
paragraph (b) of the section, failure to file an answer within the 30-
day period may be treated as a consent to the award requested.
    (b) If the FAA's counsel and the applicant believe that the issues 
in the fee application can be settled, they may jointly file a statement 
of their intent to negotiate a settlement. The filing of this statement 
shall extend the time for filing an answer for an additional 30 days, 
and further extensions may be granted by the ALJ or adjudicative officer 
upon request by the FAA's counsel and the applicant.
    (c) The answer shall explain in detail any objections to the award 
requested

[[Page 83]]

and identify the facts relied on in support of agency counsel's 
position. If the answer is based on any alleged facts not already in the 
record of the proceeding, agency counsel shall include with the answer 
either supporting affidavits or a request for further proceedings under 
Sec. 14.26.

[54 FR 46199, Nov. 1, 1989, as amended by Amdt. 14-03, 64 FR 32936, June 
18, 1999]



Sec. 14.23  Reply.

    Within 15 days after service of an answer, the applicant may file a 
reply. If the reply is based on any alleged facts not already in the 
record of the proceeding, the applicant shall include with the reply 
either supporting affidavits or a request for further proceedings under 
Sec. 14.26.



Sec. 14.24  Comments by other parties.

    Any party to a proceeding other than the applicant and the FAA's 
counsel may file comments on an application within 30 days after it is 
served, or on an answer within 15 days after it is served. A commenting 
party may not participate further in proceedings on the application 
unless the ALJ or adjudicative officer determines that the public 
interest requires such participation in order to permit full exploration 
of matters raised in the comments.

[Doc. No. FAA-1998-4379, 64 FR 32936, June 18, 1999]



Sec. 14.25  Settlement.

    The applicant and agency counsel may agree on a proposed settlement 
of the award before final action on the application, either in 
connection with a settlement of the underlying proceeding, or after the 
underlying proceeding has been concluded. If a prevailing party and 
agency counsel agree on a proposed settlement of an award before an 
application has been filed, the application shall be filed with the 
proposed settlement.



Sec. 14.26  Further proceedings.

    (a) Ordinarily the determination of an award will be made on the 
basis of the written record; however, on request of either the applicant 
or agency counsel, or on his or her own initiative, the ALJ or 
adjudicative officer assigned to the matter may order further 
proceedings, such as an informal conference, oral argument, additional 
written submissions, or an evidentiary hearing. Such further proceedings 
shall be held only when necessary for full and fair resolution of the 
issues arising from the application and shall be conducted as promptly 
as possible.
    (b) A request that the administrative law judge order further 
proceedings under this section shall specifically identify the 
information sought or the disputed issues and shall explain why the 
additional proceedings are necessary to resolve the issues.

[54 FR 46199, Nov. 1, 1989, as amended by Amdt. 14-03, 64 FR 32936, June 
18, 1999]



Sec. 14.27  Decision.

    (a) The ALJ shall issue an initial decision on the application 
within 60 days after completion of proceedings on the application.
    (b) An adjudicative officer in a proceeding under part 17 of this 
chapter and the AMS shall prepare a findings and recommendations for the 
Office of Dispute Resolution for Acquisition.
    (c) A decision under paragraph (a) or (b) of this section shall 
include written findings and conclusions on the applicant's eligibility 
and status as prevailing party and an explanation of the reasons for any 
difference between the amount requested and the amount awarded. The 
decision shall also include, if at issue, findings on whether the FAA's 
position was substantially justified, or whether special circumstances 
make an award unjust.

[Doc. No. FAA-1998-4379, 64 FR 32936, June 18, 1999]



Sec. 14.28  Review by FAA decisionmaker.

    (a) In proceedings other than those under part 17 of this chapter 
and the AMS, either the applicant or the FAA counsel may seek review of 
the initial decision on the fee application in accordance with subpart G 
of part 13 of this chapter, specifically Sec. 13.233. Additionally, the 
FAA Decisionmaker may decide to review the decision on his/her own 
initiative. If neither the applicant nor the FAA's counsel seeks review 
within 30 days after the decision is issued, it shall become final. 
Whether

[[Page 84]]

to review a decision is a matter within the discretion of the FAA 
Decisionmaker. If review is taken, the FAA Decisionmaker will issue a 
final decision on the application or remand the application to the ALJ 
who issue the initial fee award determination for further proceedings.
    (b) In proceedings under part 17 of this chapter and the AMS, the 
adjudicative officer shall prepare findings and recommendations for the 
Office of Dispute Resolution for Acquisition with recommendations as to 
whether or not an award should be made, the amount of the award, and the 
reasons therefor. The Office of Dispute Resolution for Acquisition shall 
submit a recommended order to the Administrator after the completion of 
all submissions related to the EAJA application. Upon the 
Administrator's action, the order shall become final, and may be 
reviewed under 49 U.S.C. 46110.

[Doc. No. FAA-1998-4379, 64 FR 32936, June 18, 1999, as amended at 70 FR 
8238, Feb. 18, 2005]



Sec. 14.29  Judicial review.

    If an applicant is dissatisfied with the determination of fees and 
other expenses made under this subsection, pursuant 5 U.S.C. 504(c)(2), 
that applicant may, within thirty (30) days after the determination is 
made, appeal the determination to the court of the United States having 
jurisdiction to review the merits of the underlying decision of the FAA 
adversary adjudication. The court's determination on any appeal heard 
under this paragraph shall be based solely on the factual record made 
before the FAA. The court may modify the determination of fees and other 
expenses only if the court finds that the failure to make an award of 
fees and other expenses, or the calculation of the amount of the award, 
was unsupported by substantial evidence.



Sec. 14.30  Payment of award.

    An applicant seeking payment of an award shall submit to the 
disbursing official of the FAA a copy of the FAA Decisionmaker's final 
decision granting the award, accompanied by a statement that the 
applicant will not seek review of the decision in the United States 
courts. Applications for award grants in cases involving the FAA shall 
be sent to: The Office of Accounting and Audit, AAA-1, Federal Aviation 
Administration, 800 Independence Avenue, SW., Washington, DC 20591. The 
agency will pay the amount awarded to the applicant within 60 days, 
unless judicial review of the award or of the underlying decision of the 
adversary adjudication has been sought by the applicant or any other 
party to the proceeding.



PART 15_ADMINISTRATIVE CLAIMS UNDER FEDERAL TORT CLAIMS ACT--
Table of Contents



                      Subpart A_General Procedures

Sec.
15.1 Scope of regulations.
15.3 Administrative claim, when presented; appropriate office.
15.5 Administrative claim, who may file.
15.7 Administrative claims; evidence and information to be submitted.
15.9 Investigation and examination.

Subpart B_Indemnification Under Section 1118 of the Federal Aviation Act 
                                 of 1958

15.101 Applicability.
15.103 Exclusions.
15.105 Filing of requests for indemnification.
15.107 Notification requirements.
15.109 Settlements.
15.111 Conduct of litigation.
15.113 Indemnification agreements.
15.115 Payment.

    Authority: 5 U.S.C. 301; 28 U.S.C. 2672, 2675; 49 U.S.C. 106(g), 
40113, 44721.



                      Subpart A_General Procedures

    Source: Docket No. 25264, 52 FR 18171, May 13, 1987, unless 
otherwise noted.



Sec. 15.1  Scope of regulations.

    (a) These regulations apply to claims asserted under the Federal 
Tort Claims Act, as amended, for money damages against the United States 
for injury to, or loss of property, or for personal injury or death, 
caused by the negligent or wrongful act or omission of an employee of 
the FAA acting within the scope of office or employment. The regulations 
in this part supplement the Attorney General's regulations in 28 CFR 
part 14, as amended. The regulations in 28 CFR part 14, as amended,

[[Page 85]]

and the regulations in this part apply to consideration by the FAA of 
administrative claims under the Federal Tort Claims Act.



Sec. 15.3  Administrative claim, when presented; appropriate office.

    (a) A claim is deemed to have been presented when the FAA receives, 
at a place designated in paragraph (b) of this section, an executed 
Standard Form 95 or other written notification of an incident, 
accompanied by a claim for money damages in a sum certain for injury to, 
or loss of, property or for personal injury or death, alleged to have 
occurred by reason of the incident. A claim which should have been 
presented to the FAA but which was mistakenly filed with another Federal 
agency, is deemed presented to the FAA on the date the claim is received 
by the FAA at a place designated in paragraph (b) of this section. A 
claim addressed to, or filed with, the FAA by mistake will be 
transferred to the appropriate Federal agency, if that agency can be 
determined, or returned to the claimant.
    (b) Claims shall be delivered or mailed to the Assistant Chief 
Counsel, Litigation Division, AGC-400, Federal Aviation Administration, 
800 Independence Avenue, SW., Washington, DC 20591, or alternatively, 
may be mailed or delivered to the Regional Counsel in any of the FAA 
Regional Offices or the Assistant Chief Counsel, Europe, Africa, and 
Middle East Area Office.
    (c) Claim forms are available at each location listed in paragraph 
(b) of this section.
    (d) A claim presented in accordance with this section may be amended 
by the claimant at any time prior to final FAA action or prior to the 
exercise of the claimant's option, under 28 U.S.C. 2675(a), to deem the 
agency's failure to make a final disposition of his or her claim within 
6 months after it was filed as a final denial. Each amendment to a claim 
shall be submitted in writing and signed by the claimant or the 
claimant's duly authorized agent or legal representative. Upon the 
timely filing of an amendment to a pending claim, the FAA has 6 months 
thereafter in which to make a final disposition of the claim as amended, 
and the claimant's option under 28 U.S.C. 2675(a) does not accrue until 
6 months after the filing of the amendment.

[Doc. No. 18884, 44 FR 63723, Nov. 5, 1979, as amended by Amdt. 15-1, 54 
FR 39290, Sept. 25, 1989; Amdt. 15-4, 62 FR 46866, Sept. 4, 1997]



Sec. 15.5  Administrative claim, who may file.

    (a) A claim for injury to, or loss of, property may be presented by 
the owner of the property interest which is the subject of the claim or 
by the owner's duly authorized agent or legal representative.
    (b) A claim for personal injury may be presented by the injured 
person or that person's duly authorized agent or legal representative.
    (c) A claim based on death may be presented by the executor or 
administrator of the decedent's estate or by any other person legally 
entitled to assert such a claim under applicable State law.
    (d) A claim for loss wholly compensated by an insurer with the 
rights of a subrogee may be presented by the insurer. A claim for loss 
partially compensated by an insurer with the rights of a subrogee may be 
presented by the insurer or the insured individually, as their 
respective interest appear, or jointly. Whenever an insurer presents a 
claim asserting the rights of a subrogee, it shall present with its 
claim appropriate evidence that it has the rights of a subrogee.
    (e) A claim presented by an agent or legal representative shall be 
presented in the name of the claimant, be signed by the agent or legal 
representative, show the title or legal capacity of the person signing, 
and be accompanied by evidence of authority to present a claim on behalf 
of the claimant as agent, executor, administrator, parent, guardian, or 
other representative.



Sec. 15.7  Administrative claims; evidence and information to be submitted.

    (a) Death. In support of a claim based on death, the claimant may be 
required to submit the following evidence or information:
    (1) An authenticated death certificate or other competent evidence 
showing cause of death, date of death, and age of the decedent.

[[Page 86]]

    (2) The decedent's employment or occupation at time of death, 
including monthly or yearly salary or earnings (if any), and the 
duration of last employment or occupation.
    (3) Full names, addresses, birth dates, kinship, and marital status 
of the decedent's survivors, including identification of those survivors 
who were dependent for support upon the decedent at the time of death.
    (4) Degree of support afforded by the decedent to each survivor 
dependent upon decedent for support at the time of death.
    (5) Decedent's general, physical, and mental conditions before 
death.
    (6) Itemized bills for medical and burial expenses incurred by 
reason of the incident causing death or itemized receipts of payment for 
such expenses.
    (7) If damages for pain and suffering prior to death are claimed, a 
physician's detailed statement specifying the injuries suffered, 
duration of pain and suffering, any drugs administered for pain, and the 
decedent's physical condition in the interval between injury and death.
    (8) Any other evidence or information which may have a bearing on 
either the responsibility of the United States for the death or the 
amount of damages claimed.
    (b) Personal injury. In support of a claim for personal injury, 
including pain and suffering, the claimant may be required to submit the 
following evidence or information:
    (1) A written report by the attending physician or dentist setting 
forth the nature and extent of the injuries, nature and extent of 
treatment, any degree of temporary or permanent disability, the 
prognosis, period of hospitalization, and any diminished earning 
capacity.
    (2) In addition to the report required by paragraph (b)(1) of this 
section, the claimant may be required to submit to a physical or mental 
examination by a physician employed by the FAA or another Federal 
agency. A copy of the report of the examining physician is made 
available to the claimant upon the claimant's written request if the 
claimant has, upon request, furnished the report required by paragraph 
(b)(1), and has made or agrees to make available to the FAA any other 
physician's reports previously or thereafter made on the physical or 
mental condition which is the subject matter of the claim.
    (3) Itemized bills for medical, dental, and hospital expenses 
incurred or itemized receipts of payment for such expenses.
    (4) If the prognosis reveals the necessity for future treatment, a 
statement of expected expenses for such treatment.
    (5) If a claim is made for loss of time from employment, a written 
statement from the claimant's employer showing actual time lost from 
employment, whether the claimant is a full or part-time employee, and 
wages or salary actually lost.
    (6) If a claim is made for loss of income and the claimant is self-
employed, documentary evidence showing the amount of earnings actually 
lost.
    (7) Any other evidence or information which may have a bearing on 
the responsibility of the United States for the personal injury or the 
damages claimed.
    (c) Property damage. In support of a claim for injury to or loss of 
property, real or personal, the claimant may be required to submit the 
following evidence or information:
    (1) Proof of ownership of the property interest which is the subject 
of the claim.
    (2) A detailed statement of the amount claimed with respect to each 
item of property.
    (3) An itemized receipt of payment for necessary repairs or itemized 
written estimates of the cost of such repairs.
    (4) A statement listing date of purchase, purchase price, and 
salvage value, where repair is not economical.
    (5) Any other evidence or information which may have a bearing on 
either the responsibility of the United States for the injury to or loss 
of property or the damages claimed.



Sec. 15.9  Investigation and examination.

    The FAA may investigate a claim or conduct a physical examination of 
a claimant. The FAA may request any other Federal agency to investigate 
a

[[Page 87]]

claim or conduct a physical examination of a claimant and provide a 
report of the investigation or examination to the FAA.



Subpart B_Indemnification Under Section 1118 of the Federal Aviation Act 
                                 of 1958

    Source: Amdt. 15-2, 55 FR 18710, May 3, 1990, unless otherwise 
noted.



Sec. 15.101  Applicability.

    This subpart prescribes procedural requirements for the 
indemnification of a publisher of aeronautical charts or maps under 
section 1118 of the Federal Aviation Act of 1958, as amended, when the 
publisher incurs liability as a result of publishing--
    (a) A chart or map accurately depicting a defective or deficient 
flight procedure or airway that was promulgated by the FAA; or
    (b) Aeronautical data that--
    (1) Is visually displayed in the cockpit of an aircraft; and
    (2) When visually displayed, accurately depicts a defective or 
deficient flight procedure or airway promulgated by the FAA.



Sec. 15.103  Exclusions.

    A publisher that requests indemnification under this part will not 
be indemnified if--
    (a) The complaint filed against the publisher, or demand for payment 
against the publisher, first occurred before December 19, 1985;
    (b) The publisher does not negotiate a good faith settlement;
    (c) The publisher does not conduct a good faith defense;
    (d) The defective or deficient flight procedure or airway--
    (1) Was not promulgated by the FAA;
    (2) Was not accurately depicted on the publisher's chart or map;
    (3) Was not accurately displayed on a visual display in the cockpit, 
or
    (4) Was obviously defective or deficient;
    (e) The publisher does not give notice as required by Sec. 15.107 
of this part and that failure is prejudicial to the Government; or
    (f) The publisher does not appeal a lower court's decision pursuant 
to a request by the Administrator under Sec. 15.111(d)(2) of this part.



Sec. 15.105  Filing of requests for indemnification.

    A request for indemnification under this part--
    (a) May be filed by--
    (1) A publisher described in Sec. 15.101 of this part; or
    (2) The publisher's duly authorized agent or legal representative;
    (b) Shall be filed with the Chief Counsel, Federal Aviation 
Administration, 800 Independence Avenue SW., Washington, DC 20591; and
    (c) Shall state the basis for the publisher's assertion that 
indemnification under this part is required.



Sec. 15.107  Notification requirements.

    A request for indemnification will not be considered by the FAA 
unless the following conditions are met:
    (a) The publisher must notify the Chief Counsel of the FAA, within 
the time limits prescribed in paragraph (b) or (c) of this section, of 
the publisher's first receipt of a demand for payment, or service of a 
complaint in any proceeding, federal or state, in which it appears that 
indemnification under this part may be required.
    (b) For each complaint filed, or demand for payment made, on or 
after December 19, 1985, and before June 4, 1990, the notice required by 
paragraph (a) of this section must be received by the FAA on or before 
July 2, 1990.
    (c) For each complaint filed, or demand for payment made, on or 
after June 4, 1990, the notice required by paragraph (a) of this section 
must be received by the FAA within 60 days after the day the publisher 
first receives the demand for payment or service of the complaint.
    (d) Within 5 days after the day a judgment is rendered against the 
publisher in any proceeding, or within 30 days of the denial of an 
appeal, whichever is later, the publisher must notify the FAA Chief 
Counsel that--
    (1) There is an adverse judgment against the publisher; and
    (2) The publisher has a claim for indemnification against the FAA 
arising out of that judgment.

[[Page 88]]



Sec. 15.109  Settlements.

    (a) A publisher may not settle a claim with another party, for which 
the publisher has sought, or intends to seek, indemnification under this 
part, unless--
    (1) The publisher submits a copy of the proposed settlement, and a 
statement justifying the settlement, to the Chief Counsel of the FAA; 
and
    (2) The Administrator and where necessary, the appropriate official 
of the Department of Justice, approves the proposed settlement.
    (3) The publisher submits a signed release that clearly releases the 
United States from any further liability to the publisher and the 
claimant.
    (b) If the Administrator does not approve the proposed settlement, 
the Administrator will--
    (1) So notify the publisher by registered mail within 60 days of 
receipt of the proposed settlement; and
    (2) Explain why the request for indemnification was not approved.
    (c) If the Administrator approves the proposed settlement, the 
Administrator will so notify the publisher by registered mail within 60 
days after the FAA's receipt of the proposed settlement.
    (d) If the Administrator does not have sufficient information to 
approve or disapprove the proposed settlement, the Administrator will 
request, within 60 days after receipt of the proposed settlement, the 
additional information needed to make a determination.



Sec. 15.111  Conduct of litigation.

    (a) If a lawsuit is filed against the publisher and the publisher 
has sought, or intends to seek, indemnification under this part, the 
publisher shall--
    (1) Give notice as required by Sec. 15.107 of this part;
    (2) If requested by the United States--
    (i) Implead the United States as a third-party defendant in the 
action; and
    (ii) Arrange for the removal of the action to Federal Court;
    (3) Promptly provide any additional information requested by the 
United States; and
    (4) Cooperate with the United States in the defense of the lawsuit.
    (b) If the lawsuit filed against the publisher results in a proposed 
settlement, the publisher shall submit that proposed settlement to the 
FAA for approval in accordance with Sec. 15.109 of this part.
    (c) If the lawsuit filed against the publisher results in a judgment 
against the publisher and the publisher has sought, or intends to seek, 
indemnification under this part as a result of the adverse judgment, the 
publisher shall--
    (1) Give notice to the FAA as required by Sec. 15.107(d) of this 
part;
    (2) Submit a copy of the trial court's decision to the FAA Chief 
Counsel not more than 5 business days after the adverse judgment is 
rendered; and
    (3) If an appeal is taken from the adverse judgment, submit a copy 
of the appellate decision to the FAA Chief Counsel not more than 30 days 
after that decision is rendered.
    (d) Within 60 days after receipt of the trial court's decision, the 
Administrator by registered mail will--
    (1) Notify the publisher that indemnification is required under this 
part;
    (2) Request that the publisher appeal the trial court's adverse 
decision; or
    (3) Notify the publisher that it is not entitled to indemnification 
under this part and briefly state the basis for the denial.



Sec. 15.113  Indemnification agreements.

    (a) Upon a finding of the Administrator that indemnification is 
required under this part, and after obtaining the concurrence of the 
United States Department of Justice, the FAA will promptly enter into an 
indemnification agreement providing for the payment of the costs 
specified in paragraph (c) of this section.
    (b) The indemnification agreement will be signed by the Chief 
Counsel and the publisher.
    (c) The FAA will indemnify the publisher for--
    (1) Compensatory damages awarded by the court against the publisher;
    (2) Reasonable costs and fees, including reasonable attorney fees at 
a rate not to exceed that permitted under the Equal Access to Justice 
Act (5 U.S.C. 504), and any postjudgment interest, if

[[Page 89]]

the publisher conducts a good faith defense, or pursues a good faith 
appeal, at the request, or with the concurrence, of the FAA.
    (d) Except as otherwise provided in this section, the FAA will not 
indemnify the publisher for--
    (1) Punitive or exemplary damages;
    (2) Civil or criminal fines or any other litigation sanctions;
    (3) Postjudgment interest;
    (4) Costs;
    (5) Attorney fees; or
    (6) Other incidental expenses.
    (e) The indemnification agreement must provide that the Government 
will be subrogated to all claims or rights of the publisher, including 
third-party claims, cross-claims, and counterclaims.



Sec. 15.115  Payment.

    After execution of the indemnification agreement, the FAA will 
submit the agreement to the United States Department of Justice and 
request payment, in accordance with the agreement, from the Judgment 
Fund.



PART 16_RULES OF PRACTICE FOR FEDERALLY-ASSISTED AIRPORT ENFORCEMENT
PROCEEDINGS--Table of Contents



                      Subpart A_General Provisions

Sec.
16.1 Applicability and description of part.
16.3 Definitions.
16.5 Separation of functions.

Subpart B_General Rules Applicable to Complaints, Proceedings Initiated 
                         by the FAA, and Appeals

16.11 General processes.
16.13 Filing of documents.
16.15 Service of documents on the parties and the agency.
16.17 Computation of time.
16.19 Motions.

            Subpart C_Special Rules Applicable to Complaints

16.21 Pre-complaint resolution.
16.23 Pleadings.
16.25 Dismissals.
16.26 Motions to dismiss and motions for summary judgment.
16.27 Incomplete complaints.
16.29 Investigations.
16.31 Director's determinations after investigations.
16.33 Final decisions without hearing.
16.34 Consent orders.

 Subpart D_Special Rules Applicable to Proceedings Initiated by the FAA

16.101 Basis for the initiation of agency action.
16.103 Notice of investigation.
16.105 Failure to resolve informally.

                 Subpart E_Proposed Orders of Compliance

16.109 Orders terminating eligibility for grants, cease and desist 
          orders, and other compliance orders.

                           Subpart F_Hearings

16.201 Notice and order of hearing.
16.202 Powers of a hearing officer.
16.203 Appearances, parties, and rights of parties.
16.207 Intervention and other participation.
16.209 Extension of time.
16.211 Prehearing conference.
16.213 Discovery.
16.215 Depositions.
16.217 Witnesses.
16.219 Subpoenas.
16.221 Witness fees.
16.223 Evidence.
16.225 Public disclosure of evidence.
16.227 Standard of proof.
16.229 Burden of proof.
16.231 Offer of proof.
16.233 Record.
16.235 Argument before the hearing officer.
16.237 Waiver of procedures.
16.241 Initial decisions, orders, and appeals.
16.243 Consent orders.
16.245 Associate Administrator review after a hearing.

                        Subpart G_Judicial Review

16.247 Judicial review of a final decision and order.

                    Subpart H_Ex Parte Communications

16.301 Prohibited ex parte communications.
16.303 Procedures for handling ex parte communications.
16.305 Requirement to show cause and imposition of sanction.

    Authority: 49 U.S.C. 106(g), 322, 1110, 1111, 1115, 1116, 1718(a) 
and (b), 1719, 1723, 1726, 1727, 40103(e), 40113, 40116, 44502(b), 
46101, 46104, 46110, 47104, 47106(e), 47107, 47108, 47111(d), 47122, 
47123-47125, 47133, 47151-47153, 48103.

    Source: Docket No. 27783, 61 FR 54004, October 16, 1996, unless 
otherwise noted.

[[Page 90]]



                      Subpart A_General Provisions



Sec. 16.1  Applicability and description of part.

    (a) General. The provisions of this part govern all Federal Aviation 
Administration (FAA) proceedings involving Federally-assisted airports, 
except for complaints or requests for determination filed with the 
Secretary under 14 CFR part 302, whether the proceedings are instituted 
by order of the FAA or by filing a complaint with the FAA under the 
following authorities:
    (1) 49 U.S.C. 40103(e), prohibiting the grant of exclusive rights 
for the use of any landing area or air navigation facility on which 
Federal funds have been expended (formerly section 308 of the Federal 
Aviation Act of 1958, as amended).
    (2) Requirements of the Anti-Head Tax Act, 49 U.S.C. 40116.
    (3) The assurances and other Federal obligations contained in grant-
in-aid agreements issued under the Federal Airport Act of 1946, 49 
U.S.C. 1101 et seq. (repealed 1970).
    (4) The assurances and other Federal obligations contained in grant-
in-aid agreements issued under the Airport and Airway Development Act of 
1970, as amended, 49 U.S.C. 1701 et seq.
    (5) The assurances and other Federal obligations contained in grant-
in-aid agreements issued under the Airport and Airway Improvement Act of 
1982 (AAIA), as amended and recodified, 49 U.S.C. 47101 et seq., 
specifically section 511(a), 49 U.S.C. 47107, and 49 U.S.C. 47133.
    (6) Section 505(d) of the Airport and Airway Improvement Act of 
1982, and the requirements concerning civil rights and/or Disadvantaged 
Business Enterprise (DBE) issues contained in 49 U.S.C. 47107(e) and 49 
U.S.C. 47113; 49 U.S.C. 47123; 49 U.S.C. 322, as amended; 49 CFR parts 
23 and/or 26; and/or grant assurance 30 and/or grant assurance 37.
    (7) Obligations contained in property deeds for property transferred 
pursuant to section 16 of the Federal Airport Act (49 U.S.C. 1115), 
section 23 of the Airport and Airway Development Act (49 U.S.C. 1723), 
or section 516 of the Airport and Airway Improvement Act (49 U.S.C. 
47125).
    (8) Obligations contained in property deeds for property transferred 
under the Surplus Property Act (49 U.S.C. 47151-47153).
    (b) Other agencies. Where a grant assurance concerns a statute, 
executive order, regulation, or other authority that provides an 
administrative process for the investigation or adjudication of 
complaints by a Federal agency other than the FAA, persons shall use the 
administrative process established by those authorities. Where a grant 
assurance concerns a statute, executive order, regulation, or other 
authority that enables a Federal agency other than the FAA to 
investigate, adjudicate, and enforce compliance under those authorities 
on its own initiative, the FAA may defer to that Federal agency.
    (c) Other enforcement. If a complaint or action initiated by the FAA 
involves a violation of the 49 U.S.C. subtitle VII or FAA regulations, 
except as specified in paragraphs (a)(1) and (a)(2) of this section, the 
FAA may take investigative and enforcement action under 14 CFR part 13, 
``Investigative and Enforcement Procedures.''
    (d) Effective date. This part applies to a complaint filed with the 
FAA and to an investigation initiated by the FAA on or after December 
16, 1996.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56141, Sept. 12, 2013]



Sec. 16.3  Definitions.

    Terms defined in the Acts are used as so defined. As used in this 
part:
    Act means a statute listed in Sec. 16.1 and any regulation, 
agreement, or document of conveyance issued or made under that statute.
    Administrator means the Administrator of the FAA.
    Agency means the FAA.
    Agency attorney means the Deputy Chief Counsel; the Assistant Chief 
Counsel and attorneys in the Airports/Environmental Law Division of the 
Office of the Chief Counsel; the Assistant Chief Counsel and attorneys 
in an FAA region or center who represent the FAA during the 
investigation of a complaint or at a hearing on a complaint, and who 
prosecute on behalf of the

[[Page 91]]

FAA, as appropriate. An agency attorney shall not include the Chief 
Counsel; the Assistant Chief Counsel for Litigation, or any attorney on 
the staff of the Assistant Chief Counsel for Litigation, who advises the 
Associate Administrator regarding an initial decision of the hearing 
officer or any appeal to the Associate Administrator or who is 
supervised in that action by a person who provides such advice in an 
action covered by this part.
    Agency employee means any employee of the FAA.
    Associate Administrator means the FAA Associate Administrator for 
Airports or a designee. For the purposes of this part only, Associate 
Administrator also means the Assistant Administrator for Civil Rights or 
a designee for complaints that the FAA Associate Administrator for 
Airports transfers to the Assistant Administrator for Civil Rights.
    Complainant means the person submitting a complaint.
    Complaint means a written document meeting the requirements of this 
part and filed under this part:
    (1) By a person directly and substantially affected by anything 
allegedly done or omitted to be done by any person in contravention of 
any provision of any Act, as defined in this section, as to matters 
within the jurisdiction of the Administrator, or
    (2) By a person under 49 CFR 26.105(c) against a recipient of FAA 
funds alleged to have violated a provision of 49 CFR parts 23 and/or 26.
    Decisional employee means the Administrator, Deputy Administrator, 
Associate Administrator, Director, hearing officer, or other FAA 
employee who is or who may reasonably be expected to be involved in the 
decisional process of the proceeding.
    Director means the Director of the FAA Office of Airport Compliance 
and Management Analysis, or a designee. For the purposes of this part 
only, Director also means the Deputy Assistant Administrator for Civil 
Rights for complaints that the Director of the FAA Office of Airport 
Compliance and Management Analysis transfers to the Deputy Assistant 
Administrator for Civil Rights or designee.
    Electronic filing means the process of sending electronic mail 
(email) to the FAA Part 16 Docket Clerk, with scanned documents 
attached, as a Portable Document Format (PDF) file.
    Ex parte communication means an oral or written communication not on 
the public record with respect to which reasonable prior notice to all 
parties is not given, but it shall not include requests for status 
reports on any matter or proceeding covered by this part, or 
communications between FAA employees who participate as parties to a 
hearing pursuant to 16.203(b) of this part and other parties to a 
hearing.
    Hearing officer means an attorney designated by the Deputy Chief 
Counsel in a hearing order to serve as a hearing officer in a hearing 
under this part. The following are not designated as hearing officers: 
the Chief Counsel and Deputy Chief Counsel; the Regional or Center 
Counsel and attorneys in the FAA region or center in which the 
noncompliance has allegedly occurred or is occurring; the Assistant 
Chief Counsel and attorneys in the Airports and Environmental Law 
Division of the FAA Office of the Chief Counsel; and the Assistant Chief 
Counsel and attorneys in the Litigation Division of the FAA Office of 
Chief Counsel.
    Initial decision means a decision made by the hearing officer in a 
hearing under subpart F of this part.
    Mail means U.S. first class mail; U.S. certified mail; and U.S. 
express mail. Unless otherwise noted, mail also means electronic mail 
containing PDF copies of pleadings or documents required herein.
    Noncompliance means anything done or omitted to be done by any 
person in contravention of any provision of any Act, as defined in this 
section, as to matters within the jurisdiction of the Administrator.
    Party means the complainant(s) and the respondent(s) named in the 
complaint and, after an initial determination providing an opportunity 
for hearing is issued under Sec. 16.31 and subpart E of this part, the 
agency.
    Person in addition to its meaning under 49 U.S.C. 40102(a)(33), 
includes a public agency as defined in 49 U.S.C. 47102(a)(15).

[[Page 92]]

    Personal delivery means same-day hand delivery or overnight express 
delivery service.
    Respondent means any person named in a complaint as a person 
responsible for noncompliance.
    Sponsor means:
    (1) Any public agency which, either individually or jointly with one 
or more other public agencies, has received Federal financial assistance 
for airport development or planning under the Federal Airport Act, 
Airport and Airway Development Act or Airport and Airway Improvement 
Act;
    (2) Any private owner of a public-use airport that has received 
financial assistance from the FAA for such airport; and
    (3) Any person to whom the Federal Government has conveyed property 
for airport purposes under section 13(g) of the Surplus Property Act of 
1944, as amended.
    Writing or written includes paper documents that are filed and/or 
served by mail, personal delivery, facsimile, or email (as attached PDF 
files).

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56141, Sept. 12, 2013]



Sec. 16.5  Separation of functions.

    (a) Proceedings under this part, including hearings under subpart F 
of this part, will be prosecuted by an agency attorney.
    (b) After issuance of an initial determination in which the FAA 
provides the opportunity for a hearing, an agency employee engaged in 
the performance of investigative or prosecutorial functions in a 
proceeding under this part will not, in that case or a factually related 
case, participate or give advice in an initial decision by the hearing 
officer, or a final decision by the Associate Administrator or designee 
on written appeal, and will not, except as counsel or as witness in the 
public proceedings, engage in any substantive communication regarding 
that case or a related case with the hearing officer, the Associate 
Administrator on written appeal, or agency employees advising those 
officials in that capacity.
    (c) The Chief Counsel, the Assistant Chief Counsel for Litigation, 
or an attorney on the staff of the Assistant Chief Counsel for 
Litigation advises the Associate Administrator regarding an initial 
decision, an appeal, or a final decision regarding any case brought 
under this part.



Subpart B_General Rules Applicable to Complaints, Proceedings Initiated 
                         by the FAA, and Appeals



Sec. 16.11  General processes.

    (a) Under the authority of 49 U.S.C. 40113 and 47121, the Director 
may conduct investigations, issue orders, and take such other actions as 
are necessary to fulfill the purposes of this part. This includes the 
extension of any time period prescribed, where necessary or appropriate 
for a fair and complete consideration of matters before the agency, 
prior to issuance of the Director's Determination.
    (b) Notwithstanding any other provision of this part, upon finding 
that circumstances require expedited handling of a particular case or 
controversy, the Director may issue an order directing any of the 
following prior to the issuance of the Director's Determination:
    (1) Shortening the time period for any action under this part 
consistent with due process;
    (2) If other adequate opportunity to respond to pleadings is 
available, eliminating the reply, rebuttal, or other actions prescribed 
by this part;
    (3) Designating alternative methods of service; or
    (4) Directing such other measures as may be required.
    (c) Other than those matters concerning a Corrective Action Plan, 
the jurisdiction of the Director terminates upon the issuance of the 
Director's Determination. All matters arising during the appeal period, 
such as requests for extension of time to make an appeal, will be 
addressed by the Associate Administrator.
    (d) The Director may transfer to the FAA Deputy Assistant 
Administrator for Civil Rights or Office of Civil Rights designee the 
authority to prepare and issue Director's Determinations pursuant to 
Sec. 16.31 for complaints alleging violations of section 505(d) of

[[Page 93]]

the Airport and Airway Improvement Act of 1982, and the requirements 
concerning civil rights and/or Disadvantaged Business Enterprise (DBE) 
issues contained in 49 U.S.C. 47107(e) and 49 U.S.C. 47113; 49 U.S.C. 
47123; 49 U.S.C. 322, as amended; 49 CFR parts 23 and/or 26; and/or 
grant assurance 30 and/or grant assurance 37.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56142, Sept. 12, 2013]



Sec. 16.13  Filing of documents.

    Except as otherwise provided in this part, documents shall be filed 
with the FAA during a proceeding under this part as follows:
    (a) Filing address. Documents filed under this Part shall be filed 
with the Office of the Chief Counsel, Attention: FAA Part 16 Docket 
Clerk, AGC-600, Federal Aviation Administration, 800 Independence Avenue 
SW., Washington, DC 20591. Documents to be filed with a hearing officer 
shall be filed at the address and in the manner stated in the hearing 
order.
    (b) Date and method of filing. Filing of any document shall be by 
personal delivery or mail as defined in this part, by facsimile (when 
confirmed by filing on the same date by one of the foregoing methods), 
or electronically as set forth in paragraph (h) of this section. Unless 
the date is shown to be inaccurate, documents filed with the FAA shall 
be deemed to be filed on the date of personal delivery, on the mailing 
date shown on the certificate of service, on the date shown on the 
postmark if there is no certificate of service, on the send date shown 
on the facsimile (provided filing has been confirmed through one of the 
foregoing methods), or on the mailing date shown by other evidence if 
there is no certificate of service and no postmark. Unless the date is 
shown to be inaccurate, documents filed electronically shall be deemed 
to be filed on the date shown on the certificate of service or, if none, 
the date of electronic transmission to the last party required to be 
served.
    (c) Number of copies. With the exception of electronic filing or 
unless otherwise specified, an executed original and three copies of 
each document shall be filed with the FAA Part 16 Docket Clerk. One of 
the three copies shall not be stapled, bound or hole-punched. Copies 
need not be signed, but the name of the person signing the original 
shall be shown. If a hearing order has been issued in the case, one of 
the three copies shall be filed with the hearing officer unless 
otherwise prescribed by the hearing officer.
    (d) Form. Documents filed under this part shall:
    (1) Be typewritten or legibly printed;
    (2) Include, in the case of docketed proceedings, the docket number 
of the proceeding on the front page; and
    (3) Be marked to identify personal, privileged or proprietary 
information. Decisions for the publication and release of these 
documents will be made in accordance with 5 U.S.C. 552 and 49 CFR part 
7.
    (e) Signing of documents and other papers. The original of every 
document filed shall be signed by the person filing it or the person's 
duly authorized representative. The signature shall serve as a 
certification that the signer has read the document and, based on 
reasonable inquiry and to the best of the signer's knowledge, 
information, and belief, the document is--
    (1) Consistent with this part;
    (2) Warranted by existing law or that a good faith argument exists 
for extension, modification, or reversal of existing law; and
    (3) Not interposed for any improper purpose, such as to harass or to 
cause unnecessary delay or needless increase in the cost of the 
administrative process.
    (f) Designation of person to receive service. The initial document 
filed by any person shall state on the first page the name, physical 
address, telephone number, facsimile number, if any, and email address, 
if filing electronically, of the person(s) to be served with documents 
in the proceeding. If any of these items change during the proceeding, 
the person shall promptly file notice of the change with the FAA Part 16 
Docket Clerk and the hearing officer and shall serve the notice on all 
parties.
    (g) Docket numbers. Each submission identified as a complaint under 
this part by the submitting person will be assigned a docket number.

[[Page 94]]

    (h) Electronic filing. (1) The initial complaint may be served 
electronically upon the respondent only if the respondent has previously 
agreed with the complainant in writing to participate in electronic 
filing. Documents may be filed under this Part electronically by sending 
an email containing (an) attachment(s) of (a) PDF file(s) of the 
required pleading to the FAA Docket Clerk, and the person designated in 
paragraph (h)(3) of this section.
    (2) The subject line of the email must contain the names of the 
complainant and respondent, and must contain the FAA docket number (if 
assigned). The size of each email must be less than 10 MB. Email 
attachments containing executable files (e.g., .exe and .vbs files) will 
not be accepted.
    (3) The email address at which the parties may file the documents 
described in this section is [email protected]. No 
acknowledgement or receipt will be provided by the FAA to parties using 
this method. A party filing electronically as described in this section 
must provide to the FAA Part 16 Docket Clerk and the opposing party an 
email address of the person designated by the party to receive 
pleadings.
    (4) By filing a pleading or document electronically as described in 
this section, a party waives the rights under this part for service by 
the opposing party and the FAA by methods other than email. If a party 
subsequently decides to ``opt-out'' of electronic filing, that party 
must so notify the FAA Part 16 Docket Clerk and the other party in 
writing, from which time the FAA and the parties will begin serving the 
opting-out party in accordance with Sec. Sec. 16.13 and 16.15. This 
subsection only exempts the parties from the filing and service 
requirements in Sec. 16.13(a) (with the exception that ``Documents to 
be filed with a hearing officer shall be filed at the address and in the 
manner stated in the hearing order.''), the method of filing 
requirements in Sec. 16.13(b), and the number of documents requirements 
in Sec. 16.13(c).
    (i) Internet accessibility of documents filed in the Hearing Docket. 
(1) Unless protected from public disclosure, all documents filed in the 
Hearing Docket are accessible through the Federal Docket Management 
System (FDMS): http://www.regulations.gov. To access a particular case 
file, use the FDMS number assigned to the case.
    (2) Determinations issued by the Director and Associate 
Administrator in Part 16 cases, indexes of decisions, contact 
information for the FAA Hearing Docket, the rules of practice, and other 
information are available on the FAA Office of Airports' Web site at: 
http://part16.airports.faa.gov/index.cfm.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56142, Sept. 12, 2013]



Sec. 16.15  Service of documents on the parties and the agency.

    Except as otherwise provided in this part, documents shall be served 
as follows:
    (a) Whom must be served. Copies of all documents filed with the FAA 
Part 16 Docket Clerk shall be served by the persons filing them on all 
parties to the proceeding. A certificate of service shall accompany all 
documents when they are tendered for filing and shall certify concurrent 
service on the FAA and all parties. Certificates of service shall be in 
substantially the following form:

    I hereby certify that I have this day served the foregoing [name of 
document] on the following persons at the following addresses, facsimile 
numbers (if also served by facsimile), or email address (if served 
electronically in accordance with Sec. 16.13(h)), by [specify method of 
service]:

[list persons, addresses, facsimile numbers, email addresses (as 
applicable)]

Dated this --day of --, 20--.

[signature], for [party]

    (b) Method of service. Except as otherwise agreed by the parties 
and, if applicable, the hearing officer, the method of service is the 
same as set forth in Sec. 16.13(b) for filing documents.
    (c) Where service shall be made. Service shall be made to the 
persons identified in accordance with Sec. 16.13(f). If no such person 
has been designated, service shall be made on the party.
    (d) Presumption of service. There shall be a presumption of lawful 
service--
    (1) When acknowledgment of receipt is by a person who customarily or 
in

[[Page 95]]

the ordinary course of business receives mail at the address of the 
party or of the person designated under Sec. 16.13(f);
    (2) When a properly addressed envelope, sent to the most current 
address submitted under Sec. 16.13(f), has been returned as 
undeliverable, unclaimed, or refused; or
    (3) When the party serving the document electronically has a 
confirmation statement demonstrating that the email was properly sent to 
a party correctly addressed.
    (e) Date of service. The date of service shall be determined in the 
same manner as the filing date under Sec. 16.13(b).

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56143, Sept. 12, 2013]



Sec. 16.17  Computation of time.

    This section applies to any period of time prescribed or allowed by 
this part, by notice or order of the hearing officer, or by an 
applicable statute.
    (a) The date of an act, event, or default, after which a designated 
time period begins to run, is not included in a computation of time 
under this part.
    (b) The last day of a time period is included in a computation of 
time unless it is a Saturday, Sunday, or legal holiday for the FAA, in 
which case, the time period runs until the end of the next day that is 
not a Saturday, Sunday, or legal holiday.
    (c) Whenever a party has the right or is required to do some act 
within a prescribed period after service of a document upon the party, 
and the document is served on the party by first class mail or certified 
mail, 5 days shall be added to the prescribed period.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56143, Sept. 12, 2013]



Sec. 16.19  Motions.

    (a) General. An application for an order or ruling not otherwise 
specifically provided for in this part shall be by motion. Unless 
otherwise ordered by the agency, the filing of a motion will not stay 
the date that any action is permitted or required by this part.
    (b) Form and contents. Unless made during a hearing, motions shall 
be made in writing, shall state with particularity the relief sought and 
the grounds for the relief sought, and shall be accompanied by 
affidavits or other evidence relied upon. Motions introduced during 
hearings may be made orally on the record, unless the hearing officer 
directs otherwise.
    (c) Answers to motions. Except as otherwise provided in this part, 
or except when a motion is made during a hearing, any party may file an 
answer in support of or in opposition to a motion, accompanied by 
affidavits or other evidence relied upon, provided that the answer to 
the motion is filed within 10 days after the motion has been served upon 
the person answering, or any other period set by the hearing officer. 
Where a motion is made during a hearing, the answer and the ruling 
thereon may be made at the hearing, or orally or in writing within the 
time set by the hearing officer.
    (d) Deferred actions on motions. A ruling on a motion made before 
the time set for the issuance of the Director's Determination may be 
deferred to and included with the Director's Determination.
    (e) Extension by motion. A party shall file a written motion for an 
extension of time not later than 3 business days before the document is 
due unless good cause for the late filing is shown. A party filing a 
motion for extension should attempt to obtain the concurrence of the 
opposing party. A party filing a written motion for an extension of time 
shall file the motion as required under Sec. 16.13, and serve a copy of 
the motion on all parties and the docket clerk as required under Sec. 
16.15.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56143, Sept. 12, 2013]



            Subpart C_Special Rules Applicable to Complaints



Sec. 16.21  Pre-complaint resolution.

    (a) Except for those persons filing under 49 CFR 26.105(c), prior to 
filing a complaint under this part, a person directly and substantially 
affected by the alleged noncompliance shall initiate and engage in good 
faith efforts to resolve the disputed matter informally

[[Page 96]]

with those individuals or entities believed responsible for the 
noncompliance. These efforts at informal resolution may include, without 
limitation, at the parties' expense, mediation, arbitration, or the use 
of a dispute resolution board, or other form of third party assistance. 
The FAA Airports District Office, FAA Airports Field Office, FAA 
Regional Airports Division responsible for administering financial 
assistance to the sponsor, or the FAA Office of Civil Rights will be 
available upon request to assist the parties with informal resolution.
    (b) Except for complaints filed under 49 CFR 26.105(c), a complaint 
will be dismissed under Sec. 16.27 unless the person or authorized 
representative filing the complaint certifies that:
    (1) The complainant has made substantial and reasonable good faith 
efforts to resolve the disputed matter informally prior to filing the 
complaint; and
    (2) There is no reasonable prospect for practical and timely 
resolution of the dispute.
    (c) The certification required under paragraph (b) of this section, 
shall include a brief description of the party's efforts to obtain 
informal resolution but shall not include information on monetary or 
other settlement offers made but not agreed upon in writing by all 
parties. Such efforts to resolve informally should be relatively recent 
and be demonstrated by pertinent documentation. There is no required 
form or process for informal resolution, but in each case the 
requirements to resolve the matter informally must meet the requirements 
of this paragraph.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56143, Sept. 12, 2013]



Sec. 16.23  Pleadings.

    (a) A person directly and substantially affected by any alleged 
noncompliance or a person qualified under 49 CFR 26.105(c) may file a 
complaint under this part. A person doing business with an airport and 
paying fees or rentals to the airport shall be considered directly and 
substantially affected by alleged revenue diversion as defined in 49 
U.S.C. 47107(b).
    (b) Complaints filed under this part shall--
    (1) State the name and address of each person who is the subject of 
the complaint and, with respect to each person, the specific provisions 
of each Act that the complainant believes were violated;
    (2) Include all documents then available in the exercise of 
reasonable diligence, to be offered in support of the complaint, and to 
be served upon all persons named in the complaint as persons responsible 
for the alleged action(s) or omission(s) upon which the complaint is 
based;
    (3) Provide a concise but complete statement of the facts relied 
upon to substantiate each allegation; and
    (4) Except for complaints filed under 49 CFR 26.105(c), describe how 
the complainant was directly and substantially affected by the things 
done or omitted to be done by the respondents.
    (c) Unless the complaint is dismissed pursuant to Sec. 16.25 or 
Sec. 16.27, the FAA notifies the complainant and respondent in writing 
within 20 days after the date the FAA receives the complaint that the 
complaint has been docketed.
    (d) The respondent shall file an answer within 20 days of the date 
of service of the FAA notification or, if a motion is filed under Sec. 
16.26, within 20 days of the date of service of an FAA order denying all 
or part of that motion.
    (e) The complainant may file a reply within 10 days of the date of 
service of the answer.
    (f) The respondent may file a rebuttal within 10 days of the date of 
service of the complainant's reply.
    (g) The answer, reply, and rebuttal shall, like the complaint, be 
accompanied by supporting documentation upon which the parties rely.
    (h) The answer shall deny or admit the allegations made in the 
complaint or state that the person filing the document is without 
sufficient knowledge or information to admit or deny an allegation, and 
shall assert any affirmative defense.
    (i) The answer, reply, and rebuttal shall each contain a concise but 
complete statement of the facts relied upon to substantiate the answers, 
admissions, denials, or averments made.

[[Page 97]]

    (j) Amendments or supplements to the pleadings described in this 
section will not be allowed without showing good cause through a motion 
and supporting documents.
    (k) Burden of proof. Except as used in subpart F of this part,
    (1) The burden of proof is on the complainant to show noncompliance 
with an Act or any regulation, order, agreement or document of 
conveyance issued under the authority of an Act.
    (2) Except as otherwise provided by statute or rule, the proponent 
of a motion, request, or order has the burden of proof.
    (3) A party who has asserted an affirmative defense has the burden 
of proving the affirmative defense.
    (l) Except for good cause shown through motion and supporting 
documents, discovery is not permitted except as provided in Sec. Sec. 
16.213 and 16.215.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56143, Sept. 12, 2013]



Sec. 16.25  Dismissals.

    (a) Within 20 days after the receipt of the complaint, unless a 
motion has been filed under Sec. 16.26, the Director will dismiss a 
complaint, or any claim made in a complaint, with prejudice if:
    (1) It appears on its face to be outside the jurisdiction of the 
Administrator under the Acts listed in Sec. 16.1;
    (2) On its face it does not state a claim that warrants an 
investigation or further action by the FAA; or
    (3) The complainant lacks standing to file a complaint under 
Sec. Sec. 16.3 and 16.23.
    (b) A dismissal under this section will include the reasons for the 
dismissal.

[Amdt. 16-1, 78 FR 56144, Sept. 12, 2013]



Sec. 16.26  Motions to dismiss and motions for summary judgment.

    (a) In lieu of an answer, the respondent may file a motion to 
dismiss the complaint or a motion for summary judgment on the complaint. 
The respondent may move for dismissal of the entire complaint or move 
for dismissal of particular issues from adjudication. The motion must be 
filed within 20 days after the date of service of the FAA notification 
of docketing.
    (b) Motions to dismiss. (1) A motion to dismiss shall be accompanied 
by a concise statement of the reasons for seeking dismissal. The 
respondent must show that the complaint should be dismissed, with 
prejudice, if:
    (i) It appears on its face to be outside the jurisdiction of the 
Administrator under the Acts listed in Sec. 16.1;
    (ii) On its face it does not state a claim that warrants an 
investigation or further action by the FAA; or
    (iii) The complainant lacks standing to file a complaint under 
Sec. Sec. 16.3 and 16.23.
    (2) A motion to dismiss may seek dismissal of the entire complaint 
or the dismissal of specified claims in the complaint. A motion to 
dismiss shall be accompanied by a supporting memorandum of points and 
authorities.
    (3) A complainant may file an answer to the motion to dismiss within 
10 days of the date the motion is served on the complainant, or within 
any other period set by the Director. The answer shall be accompanied by 
a concise statement of reasons for opposing dismissal, and may be 
accompanied by affidavits and other documentary evidence in support of 
that contention.
    (4) Within 30 days of the date an answer to a motion to dismiss is 
due under this section, the Director may issue an order disposing of the 
motion. If the Director denies the motion to dismiss in whole or in 
part, or grants the motion in part, then within 20 days of when the 
order is served on the respondent, the respondent shall file an answer 
to the complaint.
    (5) If the Director does not act on the motion to dismiss within 30 
days of the date an answer to a motion is due under this section, the 
respondent shall file an answer to the complaint within the next 20 
days.
    (c) Motions for summary judgment. (1) A motion for summary judgment 
may be based upon the ground that there is no genuine issue of material 
fact for adjudication and that the complaint, when viewed in the light 
most favorable to the complainant, should be summarily adjudicated in 
favor of the respondent as a matter of law. A motion for summary 
judgment may seek dismissal of the entire complaint or

[[Page 98]]

dismissal of specified claims or issues in the complaint.
    (2) The motion for summary judgment shall be accompanied by a 
concise statement of the material facts as to which the respondent 
contends there is no genuine issue of material fact. The motion may 
include affidavits and documentary evidence in support of the contention 
that there is no genuine issue of material fact in dispute.
    (3) A complainant may file an answer to the motion for summary 
judgment within 10 days of the date the motion is served on the 
complainant, or within any other period set by the Director. The answer 
shall be accompanied by a concise statement of the material facts as to 
which the complainant contends there is a genuine issue, and may be 
accompanied by affidavits and other documentary evidence in support of 
that contention.
    (4) Within 30 days of the date an answer to a motion for summary 
judgment is due under this section, the Director may issue an order 
disposing of the motion. If the Director denies the motion in whole or 
in part, or grants the motion in part, then within 20 days of when the 
order is served on the respondent, the respondent shall file an answer 
to the complaint.
    (5) If the Director does not act on the motion for summary judgment 
within 30 days of the date an answer to a motion is due under this 
section, the respondent shall file an answer to the complaint within the 
next 20 days.

[Amdt. 16-1, 78 FR 56144, Sept. 12, 2013]



Sec. 16.27  Incomplete complaints.

    (a) If a complaint is not dismissed pursuant to Sec. 16.25, but is 
deficient as to one or more of the requirements set forth in Sec. 16.21 
or Sec. 16.23(b), the Director will dismiss the complaint within 20 
days after receiving it. Dismissal will be without prejudice to the 
refiling of the complaint after amendment to correct the deficiencies. 
The Director's dismissal will include the reasons for the dismissal.
    (b) Dismissals under this section are not initial determinations, 
and appeals from decisions under this section will not be permitted.

[Amdt. 16-1, 78 FR 56144, Sept. 12, 2013]



Sec. 16.29  Investigations.

    (a) If, based on the pleadings, there appears to be a reasonable 
basis for further investigation, the FAA investigates the subject matter 
of the complaint.
    (b) The investigation may include one or more of the following, at 
the sole discretion of the FAA:
    (1) A review of the written submissions or pleadings of the parties, 
as supplemented by any informal investigation the FAA considers 
necessary and by additional information furnished by the parties at FAA 
request. In rendering its initial determination, the FAA may rely 
entirely on the complaint and the responsive pleadings provided under 
this subpart. Each party shall file documents that it considers 
sufficient to present all relevant facts and argument necessary for the 
FAA to determine whether the sponsor is in compliance.
    (2) Obtaining additional oral and documentary evidence by use of the 
agency's authority to compel production of such evidence under 49 U.S.C. 
40113 and 46104, and 49 U.S.C. 47122. The Administrator's statutory 
authority to issue compulsory process has been delegated to the Chief 
Counsel, the Deputy Chief Counsel, the Assistant Chief Counsel for 
Airports and Environmental Law, and each Assistant Chief Counsel for a 
region or center.
    (3) Conducting or requiring that a sponsor conduct an audit of 
airport financial records and transactions as provided in 49 U.S.C. 
47107 and 47121.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-1 
78 FR 56145, Sept. 12, 2013]



Sec. 16.31  Director's Determinations after investigations.

    (a) After consideration of the pleadings and other information 
obtained by the FAA after investigation, the Director will render an 
initial determination and serve it upon each party within 120 days of 
the date the last pleading specified in Sec. 16.23 was due.
    (b)(1) The Director's Determination shall include findings of fact 
and conclusions of law, accompanied by explanations and based upon all 
material issues of fact, credibility of the evidence, law and discretion 
presented on

[[Page 99]]

the record, together with a statement of the reasons therefor.
    (2) The Director shall issue a determination or rule in a party's 
favor only if the determination or ruling is in accordance with law and 
supported by a preponderance of the reliable, probative, and substantial 
evidence contained in the record.
    (c) A party adversely affected by the Director's Determination may 
appeal the initial determination as provided in Sec. 16.33. However, if 
the Director's Determination that is appealed contains a Corrective 
Action Plan, the Director has the discretion to suspend the Corrective 
Action Plan until the appeal is resolved.
    (d) If the Director's Determination finds the respondent in 
noncompliance and proposes the issuance of a compliance order, the 
initial determination will include notice of opportunity for a hearing 
under subpart F of this part if a hearing is required by statute or 
otherwise provided by the FAA. A hearing may be required by statute if 
the FAA determination would terminate eligibility for grants under 49 
U.S.C. 47114(c) or (e), or terminate payments on a grant agreement under 
49 U.S.C. subchapter 471. The respondent may elect or waive a hearing, 
as provided in subpart E of this part.
    (e) The Director will not consider requests for rehearing, 
reargument, reconsideration, or modification of a Director's 
Determination without a finding of good cause.

[Amdt. 16-1, 78 FR 56145, Sept. 12, 2013]



Sec. 16.33  Final decisions without hearing.

    (a) The Associate Administrator may transfer to the FAA Assistant 
Administrator for Civil Rights the responsibility to prepare and issue 
Final Agency Decisions pursuant to this section for appeals with issues 
concerning civil rights.
    (b) The Associate Administrator will issue a final decision on 
appeal from the Director's Determination, without a hearing, where--
    (1) The complaint is dismissed after investigation;
    (2) A hearing is not required by statute and is not otherwise made 
available by the FAA; or
    (3) The FAA provides opportunity for a hearing to the respondent and 
the respondent waives the opportunity for a hearing as provided in 
subpart E of this part.
    (c) In the cases described in paragraph (b) of this section, within 
30 days after the date of service of the initial determination, a party 
adversely affected by the Director's Determination may file in 
accordance with Sec. 16.13 and serve in accordance with Sec. 16.15 a 
simultaneous Notice of Appeal and Brief.
    (d) A reply to an appeal brief may be filed within 20 days after the 
date of service of the appeal.
    (e) On appeal, the Associate Administrator will consider the issues 
addressed in any order on a motion to dismiss or motion for summary 
judgment and any issues accepted in the Director's Determination using 
the following analysis:
    (1) Are the findings of fact each supported by a preponderance of 
reliable, probative, and substantial evidence contained in the record?
    (2) Are conclusions made in accordance with law, precedent and 
policy?
    (3) Are the questions on appeal substantial?
    (4) Have any prejudicial errors occurred?
    (f) Any new issues or evidence presented in an appeal or reply will 
not be considered unless accompanied by a petition and good cause found 
as to why the new issue or evidence was not presented to the Director. 
Such a petition must:
    (1) Set forth the new matter;
    (2) Contain affidavits of prospective witnesses, authenticated 
documents, or both, or an explanation of why such substantiation is 
unavailable; and
    (3) Contain a statement explaining why such new issue or evidence 
could not have been discovered in the exercise of due diligence prior to 
the date on which the evidentiary record closed.
    (g) The Associate Administrator will issue a final decision and 
order within 60 days after the due date of the reply.
    (h) If no appeal is filed within the time period specified in 
paragraph (c) of this section, the Director's Determination becomes the 
final decision and order of the FAA without further action. A Director's 
Determination

[[Page 100]]

that becomes final, because there is no administrative appeal, is not 
judicially reviewable.
    (i) No requests for rehearing, reargument, reconsideration, or 
modification of a final order will be considered without a finding of 
good cause.

[Amdt. 16-1, 78 FR 56145, Sept. 12, 2013]



Sec. 16.34  Consent orders.

    (a) The parties may agree at any time before the issuance of a final 
agency decision to dispose of the case by proposing a consent order. 
Good faith efforts to resolve a complaint through issuance of a consent 
order may continue throughout the administrative process. However, 
except as provided in Sec. 16.11(a), such efforts may not serve as the 
basis for extensions of the times set forth in this part.
    (b) A proposal for a consent order, specified in paragraph (a) of 
this section, shall include:
    (1) A proposed consent order;
    (2) An admission of all jurisdictional facts; and
    (3) An express waiver of the right to further procedural steps and 
of all rights of judicial review.
    (c) If the parties agree to dispose of a case by issuance of a 
consent order before the FAA issues a Director's Determination, the 
proposal for a consent order is submitted jointly by the parties to the 
Director, together with a request to adopt the consent order and dismiss 
the case. The Director may issue the consent order as an order of the 
FAA and terminate the proceeding.

[Amdt. 16-1, 78 FR 56145, Sept. 12, 2013]



 Subpart D_Special Rules Applicable to Proceedings Initiated by the FAA



Sec. 16.101  Basis for the initiation of agency action.

    The FAA may initiate its own investigation of any matter within the 
applicability of this part without having received a complaint. The 
investigation may include, without limitation, any of the actions 
described in Sec. 16.29(b).



Sec. 16.103  Notice of investigation.

    Following the initiation of an investigation under Sec. 16.101, the 
FAA sends a notice to the person(s) subject to investigation. The notice 
will set forth the areas of the agency's concern and the reasons 
therefor; request a response to the notice within 30 days of the date of 
service; and inform the respondent that the FAA will, in its discretion, 
invite good faith efforts to resolve the matter.



Sec. 16.105  Failure to resolve informally.

    If the matters addressed in the FAA notices are not resolved 
informally, the FAA may issue a Director's Determination under Sec. 
16.31.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56146, Sept. 12, 2013]



                 Subpart E_Proposed Orders of Compliance



Sec. 16.109  Orders terminating eligibility for grants, cease and
desist orders, and other compliance orders.

    (a) The agency will provide the opportunity for a hearing if, in the 
Director's determination, the agency issues or proposes to issue an 
order terminating eligibility for grants pursuant to 49 U.S.C. 47106(d), 
an order suspending the payment of grant funds pursuant to 49 U.S.C. 
47111(d); an order withholding approval of any new application to impose 
a passenger facility charge pursuant to 49 U.S.C. 47111(e); a cease and 
desist order; an order directing the refund of fees unlawfully 
collected; or any other compliance order issued by the Administrator to 
carry out the provisions of the Acts, and required to be issued after 
notice and opportunity for a hearing. In cases in which a hearing is not 
required by statute, the FAA may provide opportunity for a hearing at 
its discretion.
    (b) In a case in which the agency provides the opportunity for a 
hearing, the Director's Determination issued under Sec. 16.31 will 
include a statement of the availability of a hearing under subpart F of 
this part.
    (1) Within 20 days after service of a Director's Determination under 
Sec. 16.31 that provides an opportunity for a

[[Page 101]]

hearing a person subject to the proposed compliance order may--
    (i) Request a hearing under subpart F of this part;
    (ii) Waive hearing and appeal the Director's Determination in 
writing, as provided in Sec. 16.33;
    (iii) File, jointly with a complainant, a motion to withdraw the 
complaint and to dismiss the proposed compliance action; or
    (iv) Submit, jointly with the agency, a proposed consent order under 
Sec. 16.34(c).
    (2) If the respondent fails to file an appeal in writing within the 
time periods provided in paragraph (c) of this section, the Director's 
Determination becomes final.
    (c) The Director may either direct the respondent to submit a 
Corrective Action Plan or initiate proceedings to revoke and/or deny the 
respondent's application for Airport Improvement Program discretionary 
grants under 49 U.S.C. 47115 and general aviation airport grants under 
49 U.S.C. 47114(d) when a Director's Determination finds a respondent in 
noncompliance and does not provide for a hearing.
    (d) In the event that the respondent fails to submit, in accordance 
with a Director's Determination, a Corrective Action Plan acceptable to 
the FAA within the time provided, unless extended by the FAA for good 
cause, and/or if the respondent fails to complete the Corrective Action 
Plan as specified therein, the Director may initiate action to revoke 
and/or deny applications for Airport Improvement Program discretionary 
grants under 49 U.S.C. 47115 and general aviation airport grants under 
49 U.S.C. 47114(d).
    (e) For those violations that cannot be remedied through corrective 
action, the Director may initiate action to revoke and/or deny the 
respondent's applications for Airport Improvement Program discretionary 
grants under 49 U.S.C. 47115 and general aviation airport grants under 
49 U.S.C. 47114(d).
    (f) When the Director concludes that the respondent has fully 
complied with the Corrective Action Plan and/or when the Director 
determines that the respondent has corrected the areas of noncompliance, 
the Director will terminate the proceeding.
    (g) A complainant's standing terminates upon the issuance of a 
Director's Determination that finds a respondent in noncompliance on all 
identified issues. The complainant may not appeal the Director's 
Determination if the Director finds noncompliance on all identified 
issues.

[Amdt. 16-1, 78 FR 56146, Sept. 12, 2013]



                           Subpart F_Hearings



Sec. 16.201  Notice and order of hearing.

    (a) If a respondent is provided the opportunity for hearing in an 
initial determination and does not waive hearing, the Deputy Chief 
Counsel within 10 days after the respondent elects a hearing will issue 
and serve on the respondent and complainant a hearing order. The hearing 
order will set forth:
    (1) The allegations in the complaint, or notice of investigation, 
and the chronology and results of the investigation preliminary to the 
hearing;
    (2) The relevant statutory, judicial, regulatory, and other 
authorities;
    (3) The issues to be decided;
    (4) Such rules of procedure as may be necessary to supplement the 
provisions of this part;
    (5) The name and address of the person designated as hearing 
officer, and the assignment of authority to the hearing officer to 
conduct the hearing in accordance with the procedures set forth in this 
part; and
    (6) The date by which the hearing officer is directed to issue an 
initial decision.
    (b) Where there are no genuine issues of material fact requiring 
oral examination of witnesses, the hearing order may contain a direction 
to the hearing officer to conduct a hearing by submission of briefs and 
oral argument without the presentation of testimony or other evidence.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56146, Sept. 12, 2013]



Sec. 16.202  Powers of a hearing officer.

    In accordance with the rules of this subpart, a hearing officer may:
    (a) Give notice of, and hold, prehearing conferences and hearings;

[[Page 102]]

    (b) Administer oaths and affirmations;
    (c) Issue subpoenas authorized by law and issue notices of 
deposition requested by the parties;
    (d) Limit the frequency and extent of discovery;
    (e) Rule on offers of proof;
    (f) Receive relevant and material evidence;
    (g) Regulate the course of the hearing in accordance with the rules 
of this part to avoid unnecessary and duplicative proceedings in the 
interest of prompt and fair resolution of the matters at issue;
    (h) Hold conferences to settle or to simplify the issues by consent 
of the parties;
    (i) Dispose of procedural motions and requests;
    (j) Examine witnesses; and
    (k) Make findings of fact and conclusions of law, and issue an 
initial decision.



Sec. 16.203  Appearances, parties, and rights of parties.

    (a) Appearances. Any party may appear and be heard in person.
    (1) Any party may be accompanied, represented, or advised by an 
attorney licensed by a State, the District of Columbia, or a territory 
of the United States to practice law or appear before the courts of that 
State or territory, or by another person authorized by the hearing 
officer to be the party's representative.
    (2) An attorney, or other duly authorized representative, who 
represents a party shall file a notice of appearance in accordance with 
Sec. 16.13.
    (b) Parties and agency participation.
    (1) The parties to the hearing are the complainant(s) and 
respondent(s) named in the hearing order, and the agency. The style of 
any pleadings filed under this Subpart shall name the respondent as the 
Appellant, and the Federal Aviation Administration as the Agency.
    (2) Unless otherwise specified in the hearing order, the agency 
attorney will serve as prosecutor for the agency from the date of 
issuance of the Director's Determination providing an opportunity for 
hearing.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56146, Sept. 12, 2013]



Sec. 16.207  Intervention and other participation.

    (a) Intervention and participation by other persons are permitted 
only at the hearing stage of the complaint process and with the written 
approval of the hearing officer.
    (b) A person may submit a written motion for leave to intervene as a 
party. Except for good cause shown, a motion for leave to intervene 
shall be submitted not later than 10 days after the notice of hearing 
and hearing order.
    (c) If the hearing officer finds that intervention will not unduly 
broaden the issues or delay the proceedings and, if the person has an 
interest that will benefit the proceedings, the hearing officer may 
grant a motion for leave to intervene. The hearing officer may determine 
the extent to which an intervenor may participate in the proceedings.
    (d) Other persons may petition the hearing officer for leave to 
participate in the hearing. Participation is limited to the filing of a 
posthearing brief and reply to the hearing officer and the Associate 
Administrator. Such a brief shall be filed and served on all parties in 
the same manner as the parties' posthearing briefs are filed.
    (e) Participation under this section is at the discretion of the 
hearing officer, and no decision permitting participation shall be 
deemed to constitute an expression that the participant has such a 
substantial interest in the proceeding as would entitle it to judicial 
review of such decision.

[Amdt. 16-1, 78 FR 56146, Sept. 12, 2013]



Sec. 16.209  Extension of time.

    (a) Extension by oral agreement. The parties may agree to extend for 
a reasonable period of time for filing a document under this part. If 
the parties agree, the hearing officer shall grant one extension of time 
to each party. The party seeking the extension of time shall submit a 
draft order to the

[[Page 103]]

hearing officer to be signed by the hearing officer and filed with the 
hearing docket. The hearing officer may grant additional oral requests 
for an extension of time where the parties agree to the extension.
    (b) Extension by motion. A party shall file a written motion for an 
extension of time with the hearing officer not later than 7 days before 
the document is due unless good cause for the late filing is shown. A 
party filing a written motion for an extension of time shall serve a 
copy of the motion on each party.
    (c) Failure to rule. If the hearing officer fails to rule on a 
written motion for an extension of time by the date the document was 
due, the motion for an extension of time is deemed denied.
    (d) Effect on time limits. In a hearing required by section 519(b) 
of the Airport and Airways Improvement Act, as amended in 1987, 49 
U.S.C. 47106(e) and 47111(d), the due date for the hearing officer's 
initial decision and for the final agency decision are extended by the 
length of the extension granted by the hearing officer only if the 
hearing officer grants an extension of time as a result of an agreement 
by the parties as specified in paragraph (a) of this section or, if the 
hearing officer grants an extension of time as a result of the sponsor's 
failure to adhere to the hearing schedule. In any other hearing, an 
extension of time granted by the hearing officer for any reason extends 
the due date for the hearing officer's initial decision and for the 
final agency decision by the length of time of the hearing officer's 
decision.



Sec. 16.211  Prehearing conference.

    (a) Prehearing conference notice. The hearing officer schedules a 
prehearing conference and serves a prehearing conference notice on the 
parties promptly after being designated as a hearing officer.
    (1) The prehearing conference notice specifies the date, time, 
place, and manner (in person or by telephone) of the prehearing 
conference.
    (2) The prehearing conference notice may direct the parties to 
exchange proposed witness lists, requests for evidence and the 
production of documents in the possession of another party, responses to 
interrogatories, admissions, proposed procedural schedules, and proposed 
stipulations before the date of the prehearing conference.
    (b) The prehearing conference. The prehearing conference is 
conducted by telephone or in person, at the hearing officer's 
discretion. The prehearing conference addresses matters raised in the 
prehearing conference notice and such other matters as the hearing 
officer determines will assist in a prompt, full and fair hearing of the 
issues.
    (c) Prehearing conference report. At the close of the prehearing 
conference, the hearing officer rules on any requests for evidence and 
the production of documents in the possession of other parties, 
responses to interrogatories, and admissions; on any requests for 
depositions; on any proposed stipulations; and on any pending 
applications for subpoenas as permitted by Sec. 16.219. In addition, 
the hearing officer establishes the schedule, which shall provide for 
the issuance of an initial decision not later than 110 days after 
issuance of the Director's Determination order unless otherwise provided 
in the hearing order.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56147, Sept. 12, 2013]



Sec. 16.213  Discovery.

    (a) Discovery is limited to requests for admissions, requests for 
production of documents, interrogatories, and depositions as authorized 
by Sec. 16.215.
    (b) The hearing officer shall limit the frequency and extent of 
discovery permitted by this section if a party shows that--
    (1) The information requested is cumulative or repetitious;
    (2) The information requested may be obtained from another less 
burdensome and more convenient source;
    (3) The party requesting the information has had ample opportunity 
to obtain the information through other discovery methods permitted 
under this section; or
    (4) The method or scope of discovery requested by the party is 
unduly burdensome or expensive.

[[Page 104]]



Sec. 16.215  Depositions.

    (a) General. For good cause shown, the hearing officer may order 
that the testimony of a witness may be taken by deposition and that the 
witness produce documentary evidence in connection with such testimony. 
Generally, an order to take the deposition of a witness is entered only 
if:
    (1) The person whose deposition is to be taken would be unavailable 
at the hearing;
    (2) The deposition is deemed necessary to perpetuate the testimony 
of the witness; or
    (3) The taking of the deposition is necessary to prevent undue and 
excessive expense to a party and will not result in undue burden to 
other parties or in undue delay.
    (b) Application for deposition. Any party desiring to take the 
deposition of a witness shall make application therefor to the hearing 
officer in writing, with a copy of the application served on each party. 
The application shall include:
    (1) The name and residence of the witness;
    (2) The time and place for the taking of the proposed deposition;
    (3) The reasons why such deposition should be taken; and
    (4) A general description of the matters concerning which the 
witness will be asked to testify.
    (c) Order authorizing deposition. If good cause is shown, the 
hearing officer, in his or her discretion, issues an order authorizing 
the deposition and specifying the name of the witness to be deposed, the 
location and time of the deposition and the general scope and subject 
matter of the testimony to be taken.
    (d) Procedures for deposition. (1) Witnesses whose testimony is 
taken by deposition shall be sworn or shall affirm before any questions 
are put to them. Each question propounded shall be recorded and the 
answers of the witness transcribed verbatim.
    (2) Objections to questions or evidence shall be recorded in the 
transcript of the deposition. The interposing of an objection shall not 
relieve the witness of the obligation to answer questions, except where 
the answer would violate a privilege.
    (3) The written transcript shall be subscribed by the witness, 
unless the parties by stipulation waive the signing, or the witness is 
ill, cannot be found, or refuses to sign. The reporter shall note the 
reason for failure to sign.
    (e) Depositions of agency employees. (1) Depositions of Agency 
Employees will not be allowed except under the provisions of 49 CFR part 
9.
    (2) Such depositions will be allowed only with the specific written 
permission of the Chief Counsel or his or her designee.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56147, Sept. 12, 2013]



Sec. 16.217  Witnesses.

    (a) Each party may designate as a witness any person who is able and 
willing to give testimony that is relevant and material to the issues in 
the hearing case, subject to the limitation set forth in paragraph (b) 
of this section.
    (b) The hearing officer may exclude testimony of witnesses that 
would be irrelevant, immaterial, or unduly repetitious.
    (c) Any witness may be accompanied by counsel. Counsel representing 
a nonparty witness has no right to examine the witness or otherwise 
participate in the development of testimony.



Sec. 16.219  Subpoenas.

    (a) Request for subpoena. A party may apply to the hearing officer, 
within the time specified for such applications in the prehearing 
conference report, for a subpoena to compel testimony at a hearing or to 
require the production of documents only from the following persons:
    (1) Another party;
    (2) An officer, employee, or agent of another party;
    (3) Any other person named in the complaint as participating in or 
benefiting from the actions of the respondent alleged to have violated 
any Act;
    (4) An officer, employee, or agent of any other person named in the 
complaint as participating in or benefiting from the actions of the 
respondent alleged to have violated any Act.

[[Page 105]]

    (b) Issuance and service of subpoena. (1) The hearing officer issues 
the subpoena if the hearing officer determines that the evidence to be 
obtained by the subpoena is relevant and material to the resolution of 
the issues in the case.
    (2) Subpoenas shall be served by personal service, or upon an agent 
designated in writing for the purpose, or by certified mail, return 
receipt addressed to such person or agent. Whenever service is made by 
registered or certified mail, the date of mailing shall be considered as 
the time when service is made.
    (3) A subpoena issued under this part is effective throughout the 
United States or any territory or possession thereof.
    (c) Motions to quash or modify subpoena. (1) A party or any person 
upon whom a subpoena has been served may file a motion to quash or 
modify the subpoena with the hearing officer at or before the time 
specified in the subpoena for the filing of such motions. The applicant 
shall describe in detail the basis for the application to quash or 
modify the subpoena including, but not limited to, a statement that the 
testimony, document, or tangible evidence is not relevant to the 
proceeding, that the subpoena is not reasonably tailored to the scope of 
the proceeding, or that the subpoena is unreasonable and oppressive.
    (2) A motion to quash or modify the subpoena stays the effect of the 
subpoena pending a decision by the hearing officer on the motion.



Sec. 16.221  Witness fees.

    (a) The party on whose behalf a witness appears is responsible for 
paying any witness fees and mileage expenses.
    (b) Except for employees of the United States summoned to testify as 
to matters related to their public employment, witnesses summoned by 
subpoena shall be paid the same fees and mileage expenses as are paid to 
a witness in a court of the United States in comparable circumstances.



Sec. 16.223  Evidence.

    (a) General. A party may submit direct and rebuttal evidence in 
accordance with this section.
    (b) Requirement for written testimony and evidence. Except in the 
case of evidence obtained by subpoena, or in the case of a special 
ruling by the hearing officer to admit oral testimony, a party's direct 
and rebuttal evidence shall be submitted in written form in advance of 
the oral hearing pursuant to the schedule established in the hearing 
officer's prehearing conference report. Written direct and rebuttal fact 
testimony shall be certified by the witness as true and correct. Subject 
to the same exception (for evidence obtained by subpoena or subject to a 
special ruling by the hearing officer), oral examination of a party's 
own witness is limited to certification of the accuracy of written 
evidence, including correction and updating, if necessary, and 
reexamination following cross-examination by other parties.
    (c) Subpoenaed testimony. Testimony of witnesses appearing under 
subpoena may be obtained orally.
    (d) Cross-examination. A party may conduct cross-examination that 
may be required for disclosure of the facts, subject to control by the 
hearing officer for fairness, expedition and exclusion of extraneous 
matters.
    (e) Hearsay evidence. Hearsay evidence is admissible in proceedings 
governed by this part. The fact that evidence is hearsay goes to the 
weight of evidence and does not affect its admissibility.
    (f) Admission of evidence. The hearing officer admits evidence 
introduced by a party in support of its case in accordance with this 
section, but may exclude irrelevant, immaterial, or unduly repetitious 
evidence.
    (g) Expert or opinion witnesses. An employee of the FAA or DOT may 
not be called as an expert or opinion witness for any party other than 
the agency except as provided in Department of Transportation 
regulations at 49 CFR part 9.



Sec. 16.225  Public disclosure of evidence.

    (a) Except as provided in this section, the hearing shall be open to 
the public.
    (b) The hearing officer may order that any information contained in 
the

[[Page 106]]

record be withheld from public disclosure. Any person may object to 
disclosure of information in the record by filing a written motion to 
withhold specific information with the hearing officer. The person shall 
state specific grounds for nondisclosure in the motion.
    (c) The hearing officer shall grant the motion to withhold 
information from public disclosure if the hearing officer determines 
that disclosure would be in violation of the Privacy Act, would reveal 
trade secrets or privileged or confidential commercial or financial 
information, or is otherwise prohibited by law.



Sec. 16.227  Standard of proof.

    The hearing officer shall issue an initial decision or rule in a 
party's favor only if the decision or ruling is in accordance with law 
and supported by a preponderance of the reliable, probative, and 
substantial evidence contained in the record.

[Amdt. 16-1, as amended at 78 FR 56147, Sept. 12, 2013]



Sec. 16.229  Burden of proof.

    As used in this subpart, the burden of proof is as follows:
    (a) The burden of proof of noncompliance with an Act or any 
regulation, order, agreement or document of conveyance issued under the 
authority of an Act is on the agency.
    (b) Except as otherwise provided by statute or rule, the proponent 
of a motion, request, or order has the burden of proof.
    (c) A party who has asserted an affirmative defense has the burden 
of proving the affirmative defense.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56147, Sept. 12, 2013]



Sec. 16.231  Offer of proof.

    A party whose evidence has been excluded by a ruling of the hearing 
officer may offer the evidence on the record when filing an appeal.



Sec. 16.233  Record.

    (a) Exclusive record. The transcript of all testimony in the 
hearing, all exhibits received into evidence, all motions, applications 
requests and rulings, all documents included in the hearing record and 
the Director's Determination shall constitute the exclusive record for 
decision in the proceedings and the basis for the issuance of any 
orders.
    (b) Examination and copy of record. A copy of the record will be 
filed by the FAA Part 16 Docket Clerk in the Federal Docket Management 
System (FDMS). Any person desiring to review the record may then do so 
at http://www.regulations.gov.

[Amdt. 16-1, 78 FR 56147, Sept .12, 2013]



Sec. 16.235  Argument before the hearing officer.

    (a) Argument during the hearing. During the hearing, the hearing 
officer shall give the parties reasonable opportunity to present oral 
argument on the record supporting or opposing motions, objections, and 
rulings if the parties request an opportunity for argument. The hearing 
officer may direct written argument during the hearing if the hearing 
officer finds that submission of written arguments would not delay the 
hearing.
    (b) Posthearing briefs. The hearing officer may request or permit 
the parties to submit posthearing briefs. The hearing officer may 
provide for the filing of simultaneous reply briefs as well, if such 
filing will not unduly delay the issuance of the hearing officer's 
initial decision. Posthearing briefs shall include proposed findings of 
fact and conclusions of law; exceptions to rulings of the hearing 
officer; references to the record in support of the findings of fact; 
and supporting arguments for the proposed findings, proposed 
conclusions, and exceptions.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56147, Sept. 12, 2013]



Sec. 16.237  Waiver of procedures.

    (a) The hearing officer shall waive such procedural steps as all 
parties to the hearing agree to waive before issuance of an initial 
decision.
    (b) Consent to a waiver of any procedural step bars the raising of 
this issue on appeal.

[[Page 107]]

    (c) The parties may not by consent waive the obligation of the 
hearing officer to enter an initial decision on the record.



Sec. 16.241  Initial decisions, order, and appeals.

    (a) The hearing officer shall issue an initial decision based on the 
record developed during the proceeding and shall send the initial 
decision to the parties not later than 110 days after the Director's 
Determination unless otherwise provided in the hearing order.
    (b) Each party adversely affected by the hearing officer's initial 
decision may file an appeal with the Associate Administrator within 15 
days of the date the initial decision is issued. Each party may file a 
reply to an appeal within 10 days after it is served on the party. 
Filing and service of appeals and replies shall be by personal delivery.
    (c) If an appeal is filed, the Associate Administrator reviews the 
entire record and issues a final agency decision and order within 60 
days of the due date of the reply. If no appeal is filed, the Associate 
Administrator may take review of the case on his or her own motion. If 
the Associate Administrator finds that the respondent is not in 
compliance with any Act or any regulation, agreement, or document of 
conveyance issued or made under such Act, the final agency order 
includes, in accordance with Sec. 16.245(d), a statement of corrective 
action, if appropriate, and identifies sanctions for continued 
noncompliance.
    (d) If no appeal is filed, and the Associate Administrator does not 
take review of the initial decision on the Associate Administrator's own 
motion, the initial decision shall take effect as the final agency 
decision and order on the sixteenth day after the actual date the 
initial decision is issued.
    (e) The failure to file an appeal is deemed a waiver of any rights 
to seek judicial review of an initial decision that becomes a final 
agency decision by operation of paragraph (d) of this section.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56147, Sept. 12, 2013]



Sec. 16.243  Consent orders.

    (a) The agency attorney and the respondents may agree at any time 
before the issuance of a final decision and order to dispose of the case 
by issuance of a consent order. Good faith efforts to resolve a 
complaint through issuance of a consent order may continue throughout 
the administrative process. Except as provided in Sec. 16.209, such 
efforts may not serve as the basis for extensions of the times set forth 
in this part.
    (b) A proposal for a consent order, specified in paragraph (a) of 
this section, shall include:
    (1) A proposed consent order;
    (2) An admission of all jurisdictional facts;
    (3) An express waiver of the right to further procedural steps and 
of all rights of judicial review; and
    (4) The hearing order, if issued, and an acknowledgment that the 
hearing order may be used to construe the terms of the consent order.
    (c) If the issuance of a consent order has been agreed upon by all 
parties to the hearing, the proposed consent order shall be filed with 
the hearing officer, along with a draft order adopting the consent 
decree and dismissing the case, for the hearing officer's adoption.
    (d) The deadline for the hearing officer's initial decision and the 
final agency decision is extended by the amount of days elapsed between 
the filing of the proposed consent order with the hearing officer and 
the issuance of the hearing officer's order continuing the hearing.
    (e) If the agency attorney and sponsor agree to dispose of a case by 
issuance of a consent order before the FAA issues a hearing order, the 
proposal for a consent order is submitted jointly to the official 
authorized to issue a hearing order, together with a request to adopt 
the consent order and dismiss the case. The official authorized to issue 
the hearing order issues the consent order as an order of the FAA and 
terminates the proceeding.

[Doc. No. 27783, 61 FR 54004, October 16, 1996, as amended at Amdt. 16-
1, 78 FR 56147, Sept. 12, 2013]

[[Page 108]]



Sec. 16.245  Associate Administrator review after a hearing.

    (a) The Associate Administrator may transfer to the FAA Assistant 
Administrator for Civil Rights the authority to prepare and issue Final 
Agency Decisions pursuant to Sec. 16.241 for appeals from a hearing 
concerning civil rights issues.
    (b) After a hearing is held, and, after considering the issues as 
set forth in Sec. 16.245(e), if the Associate Administrator determines 
that the hearing officer's initial decision or order should be changed, 
the Associate Administrator may:
    (1) Make any necessary findings and issue an order in lieu of the 
hearing officer's initial decision or order, or
    (2) Remand the proceeding for any such purpose as the Associate 
Administrator may deem necessary.
    (c) If the Associate Administrator takes review of the hearing 
officer's initial decision on the Associate Administrator's own motion, 
the Associate Administrator will issue a notice of review within 20 days 
of the actual date the initial decision is issued.
    (1) The notice sets forth the specific findings of fact and 
conclusions of law in the initial decision that are subject to review by 
the Associate Administrator.
    (2) Parties may file one brief on review to the Associate 
Administrator or rely on their posthearing brief to the hearing officer. 
A brief on review shall be filed not later than 10 days after service of 
the notice of review. Filing and service of a brief on review shall be 
by personal delivery.
    (3) The Associate Administrator issues a final agency decision and 
order within 30 days of the due date of the brief. If the Associate 
Administrator finds that the respondent is not in compliance with any 
Act or any regulation, agreement or document of conveyance issued under 
such Act, the final agency order includes a statement of corrective 
action, if appropriate.
    (d) When the final agency decision finds a respondent in 
noncompliance, and where a respondent fails to properly seek judicial 
review of the final agency decision as set forth in subpart G of this 
part, the Associate Administrator will issue an order remanding the case 
to the Director for the following action:
    (1) In the event that the respondent fails to submit, in accordance 
with the final agency decision, a Corrective Action Plan acceptable to 
the FAA within the time provided, unless extended by the FAA for good 
cause, and/or if the respondent fails to complete the Corrective Action 
Plan as specified therein, the Director may initiate action to revoke 
and/or deny applications for Airport Improvement Program grants issued 
under 49 U.S.C. 47114(c)-(e) and 47115. When the Director concludes that 
the respondent has fully complied with the Corrective Action Plan, the 
Director will issue an Order terminating the proceeding.
    (2) For those violations that cannot be remedied through corrective 
action, the Director may initiate action to revoke and/or deny the 
respondent's applications for Airport Improvement Program grants issued 
under 49 U.S.C. 47114(c)-(e) and 47115.
    (e) On appeal from a hearing officer's initial decision, the 
Associate Administrator will consider the following questions:
    (1) Are the findings of fact each supported by a preponderance of 
reliable, probative and substantial evidence?
    (2) Are conclusions made in accordance with law, precedent and 
policy?
    (3) Are the questions on appeal substantial?
    (4) Have any prejudicial errors occurred?
    (f) Any new issues or evidence presented in an appeal or reply will 
not be allowed unless accompanied by a certified petition and good cause 
found as to why the new matter was not presented to the Director. Such a 
petition must:
    (1) Set forth the new issues or evidence;
    (2) Contain affidavits of prospective witnesses, authenticated 
documents, or both, or an explanation of why such substantiation is 
unavailable; and
    (3) Contain a statement explaining why such new matter could not 
have been discovered in the exercise of due diligence prior to the date 
on which the evidentiary record closed.

[[Page 109]]

    (g) A Final Agency Decision may be appealed in accordance with 
subpart G of this part.

[Amdt. 16-1, 78 FR 56147, Sept. 12, 2013]



                        Subpart G_Judicial Review

    Source: Docket No. 27783, 61 FR 54004, October 16, 1996, unless 
otherwise noted. Redesignated by Amdt. 16-1, 78 FR 56148, Sept. 12, 
2013.



Sec. 16.247  Judicial review of a final decision and order.

    (a) A person may seek judicial review, in a United States Court of 
Appeals, of a final decision and order of the Associate Administrator, 
and of an order of dismissal with prejudice issued by the Director, as 
provided in 49 U.S.C. 46110 or 49 U.S.C. 47106(d) and 47111(d). A party 
seeking judicial review shall file a petition for review with the Court 
not later than 60 days after the order has been served on the party or 
within 60 days after the entry of an order under 49 U.S.C. 46110.
    (b) The following do not constitute final decisions and orders 
subject to judicial review:
    (1) An FAA decision to dismiss a complaint without prejudice, as set 
forth in Sec. 16.27;
    (2) A Director's Determination;
    (3) An initial decision issued by a hearing officer at the 
conclusion of a hearing;
    (4) A Director's Determination or an initial decision of a hearing 
officer becomes the final decision of the Associate Administrator 
because it was not appealed within the applicable time periods provided 
under Sec. Sec. 16.33(c) and 16.241(b).

[Doc. No. 27783, 61 FR 54004, October 16, 1996. Redesignated and amended 
by Amdt. 16-1, 78 FR 56148, Sept. 12, 2013]



                    Subpart H_Ex Parte Communications

    Source: Docket No. 27783, 61 FR 54004, October 16, 1996, unless 
otherwise noted. Redesignated at 78 FR 56148, Sept. 12, 2013.



Sec. 16.301  Prohibited ex parte communications.

    (a) The prohibitions of this section shall apply from the time a 
proceeding is noticed for hearing unless the person responsible for the 
communication has knowledge that it will be noticed, in which case the 
prohibitions shall apply at the time of the acquisition of such 
knowledge.
    (b) Except to the extent required for the disposition of ex parte 
matters as authorized by law:
    (1) No interested person outside the FAA and no FAA employee 
participating as a party shall make or knowingly cause to be made to any 
decisional employee an ex parte communication relevant to the merits of 
the proceeding;
    (2) No FAA employee shall make or knowingly cause to be made to any 
interested person outside the FAA an ex parte communication relevant to 
the merits of the proceeding; or
    (3) Ex parte communications regarding solely matters of agency 
procedure or practice are not prohibited by this section.

[Doc. No. 27783, 61 FR 54004, October 16, 1996. Redesignated at Amdt. 
16-1, 78 FR 56148, Sept. 12, 2013]



Sec. 16.303  Procedures for handling ex parte communications.

    A decisional employee who receives or who makes or knowingly causes 
to be made a communication prohibited by Sec. 16.303 shall place in the 
public record of the proceeding:
    (a) All such written communications;
    (b) Memoranda stating the substance of all such oral communications; 
and
    (c) All written responses, and memoranda stating the substance of 
all oral responses, to the materials described in paragraphs (a) and (b) 
of this section.

[Doc. No. 27783, 61 FR 54004, October 16, 1996. Redesignated at Amdt. 
16-1, 78 FR 56148, Sept. 12, 2013]



Sec. 16.305  Requirement to show cause and imposition of sanction.

    (a) Upon receipt of a communication knowingly made or knowingly 
caused to be made by a party in violation of Sec. 16.303, the Associate 
Administrator or his designee or the hearing officer may,

[[Page 110]]

to the extent consistent with the interests of justice and the policy of 
the underlying statutes, require the party to show cause why his or her 
claim or interest in the proceeding should not be dismissed, denied, 
disregarded, or otherwise adversely affected on account of such 
violation.
    (b) The Associate Administrator may, to the extent consistent with 
the interests of justice and the policy of the underlying statutes 
administered by the FAA, consider a violation of this subpart sufficient 
grounds for a decision adverse to a party who has knowingly committed 
such violation or knowingly caused such violation to occur.

[Doc. No. 27783, 61 FR 54004, October 16, 1996. Redesignated at Amdt. 
16-1, 78 FR 56148, Sept. 12, 2013]



PART 17_PROCEDURES FOR PROTESTS AND CONTRACT DISPUTES--
Table of Contents



                            Subpart A_General

Sec.
17.1 Applicability.
17.3 Definitions.
17.5 Delegation of authority.
17.7 Filing and computation of time.
17.9 Protective orders.

                           Subpart B_Protests

17.11 Matters not subject to protest.
17.13 Dispute resolution process for protests.
17.15 Filing a protest.
17.17 Initial protest procedures.
17.19 Motions practice and dismissal or summary decision of protests.
17.21 Adjudicative Process for protests.
17.23 Protest remedies.

                       Subpart C_Contract Disputes

17.25 Dispute resolution process for contract disputes.
17.27 Filing a contract dispute.
17.29 Informal resolution period.
17.31 Dismissal or summary decision of contract disputes.
17.33 Adjudicative Process for contract disputes.

                Subpart D_Alternative Dispute Resolution

17.35 Use of alternative dispute resolution.
17.37 Election of alternative dispute resolution process.
17.39 Confidentiality of ADR.

                      Subpart E_Finality and Review

17.41 Final orders.
17.43 Judicial review.
17.45 Conforming amendments.
17.47 Reconsideration.

                         Subpart F_Other Matters

17.49 Sanctions.
17.51 Decorum and professional conduct.
17.53 Orders and subpoenas for testimony and document production.
17.55 Standing orders of the ODRA director.

                         Subpart G_Pre-Disputes

17.57 Dispute resolution process for Pre-disputes.
17.59 Filing a Pre-dispute.
17.61 Use of alternative dispute resolution.

Appendix A to Part 17--Alternative Dispute Resolution (ADR)

    Authority: 5 U.S.C. 570-581, 49 U.S.C. 106(f)(2), 40110, 40111, 
40112, 46102, 46014, 46105, 46109, and 46110.

    Source: 76 FR 55221, Sept. 7, 2011, unless otherwise noted.



                            Subpart A_General



Sec. 17.1  Applicability.

    This part applies to all Acquisition Management System (AMS) bid 
protests and contract disputes involving the FAA that are filed at the 
Office of Dispute Resolution for Acquisition (ODRA) on or after October 
7, 2011, with the exception of those contract disputes arising under or 
related to FAA contracts entered into prior to April 1, 1996, where such 
contracts have not been modified to be made subject to the FAA AMS. This 
part also applies to pre-disputes as described in subpart G of this 
part.



Sec. 17.3  Definitions.

    (a) Accrual means to come into existence as a legally enforceable 
claim.
    (b) Accrual of a contract claim means that all events relating to a 
claim have occurred, which fix liability of either the government or the 
contractor and permit assertion of the claim, regardless of when the 
claimant actually discovered those events. For liability to be fixed, 
some injury must have occurred. Monetary damages need not have been 
incurred, but if the claim is

[[Page 111]]

for money, such damages must be capable of reasonable estimation. The 
accrual of a claim or the running of the limitations period may be 
tolled on equitable grounds, including but not limited to active 
concealment, fraud, or if the facts were inherently unknowable.
    (c) Acquisition Management System (AMS) establishes the policies, 
guiding principles, and internal procedures for the FAA's acquisition 
system.
    (d) Adjudicative Process is an administrative adjudicatory process 
used to decide protests and contract disputes where the parties have not 
achieved resolution through informal communication or the use of ADR. 
The Adjudicative Process is conducted by a Dispute Resolution Officer 
(DRO) or Special Master selected by the ODRA Director to preside over 
the case in accordance with Public Law 108-176, Section 224, Codified at 
49 U.S.C. 40110(d)(4).
    (e) Administrator means the Administrator of the Federal Aviation 
Administration.
    (f) Alternative Dispute Resolution (ADR) is the primary means of 
voluntary dispute resolution that is employed by the ODRA. See Appendix 
A of this part.
    (g) Compensated Neutral refers to an impartial third party chosen by 
the parties to act as a facilitator, mediator, or arbitrator functioning 
to resolve the protest or contract dispute under the auspices of the 
ODRA. The parties pay equally for the services of a compensated neutral, 
unless otherwise agreed to by the parties. An ODRA DRO or neutral cannot 
be a compensated neutral.
    (h) Contract Dispute, as used in this part, means a written request 
to the ODRA seeking, as a matter of right under an FAA contract subject 
to the AMS, the payment of money in a sum certain, the adjustment or 
interpretation of contract terms, or for other relief arising under, 
relating to, or involving an alleged breach of that contract. A contract 
dispute does not require, as a prerequisite, the issuance of a 
Contracting Officer final decision. Contract disputes, for purposes of 
ADR only, may also involve contracts not subject to the AMS.
    (i) Counsel refers to a Legal Representative who is an attorney 
licensed by a State, the District of Columbia, or a territory of the 
United States to practice law or appear before the courts of that State 
or territory.
    (j) Contractor is a party in contractual privity with the FAA and 
responsible for performance of a contract's requirements.
    (k) Discovery is the procedure whereby opposing parties in a protest 
or contract dispute may, either voluntarily or to the extent ordered by 
the ODRA, obtain testimony from, or documents and information held by, 
other parties or non-parties.
    (l) Dispute Resolution Officer (DRO) is an attorney and member of 
the ODRA staff. The term DRO can include the Director of the ODRA.
    (m) Interested party, in the context of a bid protest, is one whose 
direct economic interest has been or would be affected by the award or 
failure to award an FAA contract. Proposed subcontractors are not 
``interested parties'' within this definition and are not eligible to 
submit protests to the ODRA. Subcontractors not in privity with the FAA 
are not interested parties in the context of a contract dispute.
    (n) Intervenor is an interested party other than the protester whose 
participation in a protest is allowed by the ODRA. For a post-award 
protest, the awardee of the contract that is the subject of the protest 
will be allowed, upon timely request, to participate as an intervenor in 
the protest. In such a protest, no other interested parties will be 
allowed to participate as intervenors.
    (o) Legal Representative is an individual(s) designated to act on 
behalf of a party in matters before the ODRA. Unless otherwise provided 
under Sec. Sec. 17.15(c)(2), 17.27(a)(1), or 17.59(a)(6), a Notice of 
Appearance must be filed with the ODRA containing the name, address, 
telephone and facsimile (Fax) numbers of a party's legal representative.
    (p) Neutral refers to an impartial third party in the ADR process 
chosen by the parties to act as a facilitator, mediator, arbitrator, or 
otherwise to aid the parties in resolving a protest or contract dispute. 
A neutral can be a

[[Page 112]]

DRO or a person not an employee of the ODRA.
    (q) ODRA is the FAA's exclusive forum acting on behalf of the 
Administrator, pursuant to the statutory authority granted by Public Law 
108-176, Section 224, to provide dispute resolution services and to 
adjudicate matters within its jurisdiction. The ODRA may also provide 
non-binding dispute resolution services in matters outside of its 
jurisdiction where mutually requested to do so by the parties involved.
    (r) Parties include the protester(s) or the contractor, the FAA, and 
any intervenor(s).
    (s) Pre-Disputes mean an issue(s) in controversy concerning an FAA 
contract or solicitation that, by mutual agreement of the parties, is 
filed with the ODRA. See subpart G of this part.
    (t) Product Team, as used in these rules, refers to the FAA 
organization(s) responsible for the procurement or contracting activity, 
without regard to funding source, and includes the Contracting Officer 
(CO). The Product Team, acting through assigned FAA counsel, is 
responsible for all communications with and submissions to the ODRA in 
pending matters.
    (u) Screening Information Request (SIR or Solicitation) means a 
request by the FAA for documentation, information, presentations, 
proposals, or binding offers concerning an approach to meeting potential 
acquisition requirements established by the FAA.
    (v) A Special Master is a non-FAA attorney or judge who has been 
assigned by the ODRA to act as its finder of fact, and to make findings 
and recommendations based upon AMS policy and applicable law and 
authorities in the Adjudicative Process.



Sec. 17.5  Delegation of authority.

    (a) The authority of the Administrator to conduct dispute resolution 
and adjudicative proceedings concerning acquisition matters is delegated 
to the Director of the ODRA.
    (b) The Director of the ODRA may redelegate to Special Masters and 
DROs such delegated authority in paragraph (a) of this section as deemed 
necessary by the Director for efficient resolution of an assigned 
protest or contract dispute, including the imposition of sanctions for 
the filing of frivolous pleadings, making false statements, or other 
disciplinary actions. See subpart F of this part.



Sec. 17.7  Filing and computation of time.

    (a) Filing of a protest or contract dispute may be accomplished by 
overnight delivery, by hand delivery, by Fax, or, if permitted by Order 
of the ODRA, by electronic filing. A protest or contract dispute is 
considered to be filed on the date it is received by the ODRA during 
normal business hours. The ODRA's normal business hours are from 8:30 
a.m. to 5 p.m. Eastern Time. A protest or contract dispute received 
after the time period prescribed for filing shall not be considered 
timely filed. Service shall also be made on the Contracting Officer (CO) 
pursuant to Sec. Sec. 17.15(e) and 17.27(d).
    (b) Submissions to the ODRA after the initial filing of a protest or 
contract dispute may be accomplished by any means available in paragraph 
(a) of this section. Copies of all such submissions shall be served on 
the opposing party or parties.
    (c) The time limits stated in this part are calculated in business 
days, which exclude weekends, Federal holidays and other days on which 
Federal Government offices in Washington, DC are not open. In computing 
time, the day of the event beginning a period of time shall not be 
included. If the last day of a period falls on a weekend or a Federal 
holiday, the first business day following the weekend or holiday shall 
be considered the last day of the period.
    (d) Electronic Filing--Procedures for electronic filing may be 
utilized where permitted by Order of the ODRA on a case-by-case basis or 
pursuant to a Standing Order of the ODRA permitting electronic filing.



Sec. 17.9  Protective orders.

    (a) The ODRA may issue protective orders addressing the treatment of 
protected information, including protected information in electronic 
form, either at the request of a party or upon its own initiative. Such 
information may include proprietary, confidential, or source-selection-
sensitive material, or other information the release of

[[Page 113]]

which could result in a competitive advantage to one or more firms.
    (b) The terms of the ODRA's standard protective order may be altered 
to suit particular circumstances, by negotiation of the parties, subject 
to the approval of the ODRA. The protective order establishes procedures 
for application for access to protected information, identification and 
safeguarding of that information, and submission of redacted copies of 
documents omitting protected information.
    (c) After a protective order has been issued, counsel or consultants 
retained by counsel appearing on behalf of a party may apply for access 
to the material under the order by submitting an application to the 
ODRA, with copies furnished simultaneously to all parties. The 
application shall establish that the applicant is not involved in 
competitive decision-making for any firm that could gain a competitive 
advantage from access to the protected information and that the 
applicant will diligently protect any protected information received 
from inadvertent disclosure. Objections to an applicant's admission 
shall be raised within two (2) days of the application, although the 
ODRA may consider objections raised after that time for good cause.
    (d) Any violation of the terms of a protective order may result in 
the imposition of sanctions, including but not limited to removal of the 
violator from the protective order and reporting of the violator to his 
or her bar association(s), and the taking of other actions as the ODRA 
deems appropriate. Additional civil or criminal penalties may apply.



                           Subpart B_Protests



Sec. 17.11  Matters not subject to protest.

    The following matters may not be protested before the ODRA, except 
for review of compliance with the AMS:
    (a) FAA purchases from or through, State, local, and tribal 
governments and public authorities;
    (b) FAA purchases from or through other Federal agencies;
    (c) Grants;
    (d) Cooperative agreements;
    (e) Other transactions.



Sec. 17.13  Dispute resolution process for protests.

    (a) Protests concerning FAA SIRs, solicitations, or contract awards 
shall be resolved pursuant to this part.
    (b) Potential protestors should, where possible, attempt to resolve 
any issues concerning potential protests with the CO. Such attempts are 
not a prerequisite to filing a protest with the ODRA.
    (c) Offerors or prospective offerors shall file a protest with the 
ODRA in accordance with Sec. 17.15. The protest time limitations set 
forth in Sec. 17.15 will not be extended by attempts to resolve a 
potential protest with the CO. Other than the time limitations specified 
in Sec. 17.15 for the filing of protests, the ODRA retains the 
discretion to modify any timeframes established herein in connection 
with protests.
    (d) In accordance with Sec. 17.17(b), the ODRA shall convene an 
initial status conference for the purpose of scheduling proceedings in 
the protest and to encourage the parties to consider using the ODRA's 
ADR process to attempt to resolve the protest, pursuant to subpart D of 
this part. It is the Agency's policy to use voluntary ADR to the maximum 
extent practicable. If the parties elect not to attempt ADR, or if ADR 
efforts do not completely resolve the protest, the protest will proceed 
under the ODRA Adjudicative Process set forth in subpart E of this part. 
Informal ADR techniques may be utilized simultaneously with ongoing 
adjudication.
    (e) The ODRA Director shall designate DROs, outside neutrals or 
Special Masters as potential neutrals for the resolution of protests 
through ADR. The ultimate choice of an ADR neutral is made by the 
parties participating in the ADR. The ODRA Director also shall, at his 
or her sole discretion, designate an adjudicating DRO or Special Master 
for each matter. A person serving as a neutral in an ADR effort in a 
matter, shall not serve as an adjudicating DRO or Special Master for 
that matter.
    (f) Multiple protests concerning the same SIR, solicitation, or 
contract

[[Page 114]]

award may be consolidated at the discretion of the ODRA Director, and 
assigned to a single DRO or Special Master for adjudication.
    (g) Procurement activities, and, where applicable, contractor 
performance pending resolution of a protest, shall continue during the 
pendency of a protest, unless there is a compelling reason to suspend 
all or part of the procurement activities or contractor performance. 
Pursuant to Sec. Sec. 17.15(d) and 17.17(a), the ODRA may impose a 
temporary suspension and recommend suspension of award or contract 
performance, in whole or in part, for a compelling reason. A decision to 
suspend procurement activities or contractor performance is made in 
writing by the Administrator or the Administrator's delegee upon 
recommendation of the ODRA.



Sec. 17.15  Filing a protest.

    (a) An interested party may initiate a protest by filing with the 
ODRA in accordance with Sec. 17.7(a) within the timeframes set forth in 
this Section. Protests that are not timely filed shall be dismissed. The 
timeframes applicable to the filing of protests are as follows:
    (1) Protests based upon alleged SIR or solicitation improprieties 
that are apparent prior to bid opening or the time set for receipt of 
initial proposals shall be filed prior to bid opening or the time set 
for the receipt of initial proposals.
    (2) In procurements where proposals are requested, alleged 
improprieties that do not exist in the initial solicitation, but which 
are subsequently incorporated into the solicitation, must be protested 
not later than the next closing time for receipt of proposals following 
the incorporation.
    (3) For protests other than those related to alleged solicitation 
improprieties, the protest must be filed on the later of the following 
two dates:
    (i) Not later than seven (7) business days after the date the 
protester knew or should have known of the grounds for the protest; or
    (ii) If the protester has requested a post-award debriefing from the 
FAA Product Team, not later than five (5) business days after the date 
on which the Product Team holds that debriefing.
    (b) Protests shall be filed at:
    (1) ODRA, AGC-70, Federal Aviation Administration, 800 Independence 
Avenue, SW., Room 323, Washington, DC 20591; Telephone: (202) 267-3290, 
Fax: (202) 267-3720; or
    (2) Other address as shall be published from time to time in the 
Federal Register.
    (c) A protest shall be in writing, and set forth:
    (1) The protester's name, address, telephone number, and FAX number;
    (2) The name, address, telephone number, and FAX number of the 
protester's legal representative, and who shall be duly authorized to 
represent the protester, to be the point of contact;
    (3) The SIR number or, if available, the contract number and the 
name of the CO;
    (4) The basis for the protester's status as an interested party;
    (5) The facts supporting the timeliness of the protest;
    (6) Whether the protester requests a protective order, the material 
to be protected, and attach a redacted copy of that material;
    (7) A detailed statement of both the legal and factual grounds of 
the protest, and one (1) copy of each relevant document;
    (8) The remedy or remedies sought by the protester, as set forth in 
Sec. 17.23;
    (9) The signature of the legal representative, or another person 
duly authorized to represent the protester.
    (d) If the protester wishes to request a suspension of the 
procurement or contract performance, in whole or in part, and believes 
that a compelling reason(s) exists to suspend the procurement or 
contract performance because of the protested action, the protester 
shall, in its initial filing:
    (1) Set forth such compelling reason(s), supply all facts and 
documents supporting the protester's position; and
    (2) Demonstrate--
    (i) The protester has alleged a substantial case;
    (ii) The lack of a suspension would be likely to cause irreparable 
injury;
    (iii) The relative hardships on the parties favor a suspension; and

[[Page 115]]

    (iv) That a suspension is in the public interest.
    (3) Failure of a protester to provide information or documents in 
support of a requested suspension or failure to address the elements of 
paragraph (d)(2) of this section may result in the summary rejection of 
the request for suspension, or a requirement that the protester 
supplement its request prior to the scheduling of a Product Team 
response to the request under Sec. 17.17(a).
    (e) Concurrent with the filing of a protest with the ODRA, the 
protester shall serve a copy of the protest on the CO and any other 
official designated in the SIR for receipt of protests, by means 
reasonably calculated to be received by the CO on the same day as it is 
to be received by the ODRA. The protest shall include a signed statement 
from the protester, certifying to the ODRA the manner of service, date, 
and time when a copy of the protest was served on the CO and other 
designated official(s).
    (f) Upon receipt of the protest, the CO shall notify the awardee of 
a challenged contract award in writing of the existence of the protest. 
The awardee and/or interested parties shall notify the ODRA in writing, 
of their interest in participating in the protest as intervenors within 
two (2) business days of receipt of the CO's notification, and shall, in 
such notice, designate a person as the point of contact for the ODRA.
    (g) The ODRA has discretion to designate the parties who shall 
participate in the protest as intervenors. In protests of awarded 
contracts, only the awardee may participate as an intervenor as a matter 
of right.



Sec. 17.17  Initial protest procedures.

    (a) If, as part of its initial protest filing, the protester 
requests a suspension of procurement activities or contractor 
performance in whole or in part, in accordance with Sec. 17.15(d), the 
Product Team shall submit a response to the request to the ODRA by no 
later than the close of business on the date of the initial scheduling 
conference or on such other date as is established by the ODRA. Copies 
of the response shall be furnished to the protester and any 
intervenor(s) so as to be received within the same timeframe. The 
protester and any intervenor(s) shall have the opportunity of providing 
additional comments on the response within two (2) business days of 
receiving it. Based on its review of such submissions, the ODRA, in its 
discretion, may--
    (1) Decline the suspension request; or
    (2) Recommend such suspension to the Administrator or the 
Administrator's designee. The ODRA also may impose a temporary 
suspension of no more than ten (10) business days, where it is 
recommending that the Administrator impose a suspension.
    (b) Within five (5) business days of the filing of a protest, or as 
soon thereafter as practicable, the ODRA shall convene an initial status 
conference for purposes of:
    (1) Reviewing the ODRA's ADR and adjudication procedures and 
establishing a preliminary schedule;
    (2) Identifying legal or other preliminary or potentially 
dispositive issues and answering the parties' questions regarding the 
ODRA process;
    (3) Dealing with issues related to protected information and the 
issuance of any needed protective order;
    (4) Encouraging the parties to consider using ADR;
    (5) Appointing a DRO as a potential ADR neutral to assist the 
parties in considering ADR options and developing an ADR agreement; and
    (6) For any other reason deemed appropriate by the DRO or by the 
ODRA.
    (c) The Product Team and protester will have five (5) business days 
from the date of the initial status conference to decide whether they 
will attempt to use an ADR process in the case. With the agreement of 
the ODRA, ADR may be used concurrently with the adjudication of a 
protest. See Sec. 17.37(e).
    (d) If the Product Team and protester elect to use ADR proceedings 
to resolve the protest, they will agree upon the neutral to conduct the 
ADR proceedings (either an ODRA DRO or a compensated neutral of their 
own choosing) pursuant to Sec. 17.37, and shall execute and file with 
the ODRA a written ADR agreement. Agreement of any intervenor(s) to the 
use of ADR or the resolution of a dispute through ADR shall not be 
required.

[[Page 116]]

    (e) If the Product Team or protester indicate that ADR proceedings 
will not be used, or if ADR is not successful in resolving the entire 
protest, the ODRA Director upon being informed of the situation, will 
schedule an adjudication of the protest.



Sec. 17.19  Motions practice and dismissal or summary decision of protests.

    (a) Separate motions generally are discouraged in ODRA bid protests. 
Counsel and parties are encouraged to incorporate any such motions in 
their respective agency responses or comments. Parties and counsel are 
encouraged to attempt to resolve typical motions issues through the ODRA 
ADR process. The ODRA may rule on any non-dispositive motion, where 
appropriate and necessary, after providing an opportunity for briefing 
on the motion by all affected parties. Unjustifiable, inappropriate use 
of motions may result in the imposition of sanctions. Where appropriate, 
a party may request by dispositive motion to the ODRA, or the ODRA may 
recommend or order, that:
    (1) The protest, or any count or portion of a protest, be dismissed 
for lack of jurisdiction, timeliness, or standing to pursue the protest;
    (2) The protest, or any count or portion of a protest, be dismissed, 
if frivolous or without basis in fact or law, or for failure to state a 
claim upon which relief may be had;
    (3) A summary decision be issued with respect to the protest, or any 
count or portion of a protest, if there are no material facts in dispute 
and a party is entitled to summary decision as a matter of law.
    (b) In connection with consideration of possible dismissal or 
summary decision, the ODRA shall consider any material facts in dispute, 
in a light most favorable to the party against whom the dismissal or 
summary decision would operate and draw all factual inferences in favor 
of the non-moving party.
    (c) Either upon motion by a party or on its own initiative, the ODRA 
may, at any time, exercise its discretion to:
    (1) Recommend to the Administrator dismissal or the issuance of a 
summary decision with respect to the entire protest;
    (2) Dismiss the entire protest or issue a summary decision with 
respect to the entire protest, if delegated that authority by the 
Administrator; or
    (3) Dismiss or issue a summary decision with respect to any count or 
portion of a protest.
    (d) A dismissal or summary decision regarding the entire protest by 
either the Administrator, or the ODRA by delegation, shall be construed 
as a final agency order. A dismissal or summary decision that does not 
resolve all counts or portions of a protest shall not constitute a final 
agency order, unless and until such dismissal or decision is 
incorporated or otherwise adopted in a decision by the Administrator (or 
the ODRA, by delegation) regarding the entire protest.
    (e) Prior to recommending or entering either a dismissal or a 
summary decision, either in whole or in part, the ODRA shall afford all 
parties against whom the dismissal or summary decision is to be entered 
the opportunity to respond to the proposed dismissal or summary 
decision.



Sec. 17.21  Adjudicative Process for protests.

    (a) Other than for the resolution of preliminary or dispositive 
matters, the Adjudicative Process for protests will be commenced by the 
ODRA Director pursuant to Sec. 17.17(e).
    (b) The Director of the ODRA shall appoint a DRO or a Special Master 
to conduct the adjudication proceedings, develop the administrative 
record, and prepare findings and recommendations for review of the ODRA 
Director.
    (c) The DRO or Special Master may conduct such proceedings and 
prepare procedural orders for the proceedings as deemed appropriate; and 
may require additional submissions from the parties.
    (d) The Product Team response to the protest will be due to be filed 
and served ten (10) business days from the commencement of the ODRA 
Adjudication process. The Product Team response shall consist of a 
written chronological, supported statement of proposed facts, and a 
written presentation of applicable legal or other defenses.

[[Page 117]]

The Product Team response shall cite to and be accompanied by all 
relevant documents, which shall be chronologically indexed, individually 
tabbed, and certified as authentic and complete. A copy of the response 
shall be furnished so as to be received by the protester and any 
intervenor(s) on the same date it is filed with the ODRA. In all cases, 
the Product Team shall indicate the method of service used.
    (e) Comments of the protester and the intervenor on the Product Team 
response will be due to be filed and served five (5) business days after 
their receipt of the response. Copies of such comments shall be provided 
to the other participating parties by the same means and on the same 
date as they are furnished to the ODRA. Comments may include any 
supplemental relevant documents.
    (f) The ODRA may alter the schedule for filing of the Product Team 
response and the comments for good cause or to accommodate the 
circumstances of a particular protest.
    (g) The DRO or Special Master may convene the parties and/or their 
representatives, as needed for the Adjudicative Process.
    (h) If, in the sole judgment of the DRO or Special Master, the 
parties have presented written material sufficient to allow the protest 
to be decided on the record presented, the DRO or Special Master shall 
have the discretion to decide the protest on that basis.
    (i) The parties may engage in limited, focused discovery with one 
another and, if justified, with non-parties, so as to obtain information 
relevant to the allegations of the protest.
    (1) The DRO or Special Master shall manage the discovery process, 
including limiting its length and availability, and shall establish 
schedules and deadlines for discovery, which are consistent with 
timeframes established in this part and with the FAA policy of providing 
fair and expeditious dispute resolution.
    (2) The DRO or Special Master may also direct the parties to 
exchange, in an expedited manner, relevant, non-privileged documents.
    (3) Where justified, the DRO or Special Master may direct the taking 
of deposition testimony, however, the FAA dispute resolution process 
does not contemplate extensive discovery.
    (4) The use of interrogatories and requests for admission is not 
permitted in ODRA bid protests.
    (5) Where parties cannot voluntarily reach agreement on a discovery-
related issue, they may timely seek assistance from an ODRA ADR neutral 
or may file an appropriate motion with the ODRA. Parties may request a 
subpoena.
    (6) Discovery requests and responses are not part of the record and 
will not be filed with the ODRA, except in connection with a motion or 
other permissible filing.
    (7) Unless timely objection is made, documents properly filed with 
the ODRA will be deemed admitted into the administrative record.
    (j) Hearings are not typically held in bid protests. The DRO or 
Special Master may conduct hearings, and may limit the hearings to the 
testimony of specific witnesses and/or presentations regarding specific 
issues. The DRO or Special Master shall control the nature and conduct 
of all hearings, including the sequence and extent of any testimony. 
Hearings will be conducted:
    (1) Where the DRO or Special Master determines that there are 
complex factual issues in dispute that cannot adequately or efficiently 
be developed solely by means of written presentations and/or that 
resolution of the controversy will be dependent on his/her assessment of 
the credibility of statements provided by individuals with first-hand 
knowledge of the facts; or
    (2) Upon request of any party to the protest, unless the DRO or 
Special Master finds specifically that a hearing is unnecessary and that 
no party will be prejudiced by limiting the record in the adjudication 
to the parties' written submissions. All witnesses at any such hearing 
shall be subject to cross-examination by the opposing party and to 
questioning by the DRO or Special Master.
    (k) The Director of the ODRA may review the status of any protest in 
the Adjudicative Process with the DRO or Special Master.
    (l) After the closing of the administrative record, the DRO or 
Special

[[Page 118]]

Master will prepare and submit findings and recommendations to the ODRA 
that shall contain the following:
    (1) Findings of fact;
    (2) Application of the principles of the AMS, and any applicable law 
or authority to the findings of fact;
    (3) A recommendation for a final FAA order; and
    (4) If appropriate, suggestions for future FAA action.
    (m) In preparing findings and recommendations in protests, the DRO 
or Special Master, using the preponderance of the evidence standard, 
shall consider whether the Product Team actions in question were 
consistent with the requirements of the AMS, had a rational basis, and 
whether the Product Team decision was arbitrary, capricious or an abuse 
of discretion. Notwithstanding the above, allegations that government 
officials acted with bias or in bad faith must be established by clear 
and convincing evidence.
    (n) The DRO or Special Master has broad discretion to recommend a 
remedy that is consistent with Sec. 17.23.
    (o) A DRO or Special Master shall submit findings and 
recommendations only to the Director of the ODRA or the Director's 
designee. The findings and recommendations will be released to the 
parties and to the public upon issuance of the final FAA order in the 
case. If an ODRA protective order was issued in connection with the 
protest, or if a protest involves proprietary or competition-sensitive 
information, a redacted version of the findings and recommendations, 
omitting any protected information, shall be prepared wherever possible 
and released to the public, as soon as is practicable, along with a copy 
of the final FAA order. Only persons admitted by the ODRA under the 
protective order and Government personnel shall be provided copies of 
the unredacted findings and recommendations that contain proprietary or 
competition-sensitive information.
    (p) Other than communications regarding purely procedural matters or 
ADR, there shall be no substantive ex parte communication between ODRA 
personnel and any principal or representative of a party concerning a 
pending or potentially pending matter. A potential or serving ADR 
neutral may communicate on an ex parte basis to establish or conduct the 
ADR.



Sec. 17.23  Protest remedies.

    (a) The ODRA has broad discretion to recommend and impose protest 
remedies that are consistent with the AMS and applicable law. Such 
remedies may include, but are not limited to one or a combination of, 
the following:
    (1) Amend the SIR;
    (2) Refrain from exercising options under the contract;
    (3) Issue a new SIR;
    (4) Require a recompetition or revaluation;
    (5) Terminate an existing contract for the FAA's convenience;
    (6) Direct an award to the protester;
    (7) Award bid and proposal costs; or
    (8) Any other remedy consistent with the AMS that is appropriate 
under the circumstances.
    (b) In determining the appropriate recommendation, the ODRA may 
consider the circumstances surrounding the procurement or proposed 
procurement including, but not limited to: the nature of the procurement 
deficiency; the degree of prejudice to other parties or to the integrity 
of the acquisition system; the good faith of the parties; the extent of 
performance completed; the feasibility of any proposed remedy; the 
urgency of the procurement; the cost and impact of the recommended 
remedy; and the impact on the Agency's mission.
    (c) Attorney's fees of a prevailing protester are allowable to the 
extent permitted by the Equal Access to Justice Act, 5 U.S.C. 504(a)(1) 
(EAJA) and 14 CFR part 14.



                       Subpart C_Contract Disputes



Sec. 17.25  Dispute resolution process for contract disputes.

    (a) All contract disputes arising under contracts subject to the AMS 
shall be resolved under this subpart.
    (b) Contract disputes shall be filed with the ODRA pursuant to Sec. 
17.27.
    (c) The ODRA has broad discretion to recommend remedies for a 
contract dispute that are consistent with the AMS and applicable law, 
including

[[Page 119]]

such equitable remedies or other remedies as it deems appropriate.



Sec. 17.27  Filing a contract dispute.

    (a) Contract disputes must be in writing and should contain:
    (1) The contractor's name, address, telephone and Fax numbers and 
the name, address, telephone and Fax numbers of the contractor's legal 
representative(s) (if any) for the contract dispute;
    (2) The contract number and the name of the Contracting Officer;
    (3) A detailed chronological statement of the facts and of the legal 
grounds underlying the contract dispute, broken down by individual claim 
item, citing to relevant contract provisions and attaching copies of the 
contract and other relevant documents;
    (4) Information establishing the ODRA's jurisdiction and the 
timeliness of the contract dispute;
    (5) A request for a specific remedy, and the amount, if known, of 
any monetary remedy requested, together with pertinent cost information 
and documentation (e.g., invoices and cancelled checks). Supporting 
documentation should be broken down by individual claim item and 
summarized; and
    (6) The signature of a duly authorized representative of the 
initiating party.
    (b) Contract disputes shall be filed at the following address: ODRA, 
AGC-70, Federal Aviation Administration, 800 Independence Avenue, SW., 
Room 323, Washington, DC 20591; Telephone: (202) 267-3290, Fax: (202) 
267-3720.
    (c) A contract dispute against the FAA shall be filed with the ODRA 
within two (2) years of the accrual of the contract claim involved. A 
contract dispute by the FAA against a contractor (excluding contract 
disputes alleging warranty issues, fraud or latent defects) likewise 
shall be filed within two (2) years of the accrual of the contract 
claim. If an underlying contract entered into prior to the effective 
date of this part provides for time limitations for filing of contract 
disputes with the ODRA, which differ from the aforesaid two (2) year 
period, the limitation periods in the contract shall control over the 
limitation period of this section. In no event will either party be 
permitted to file with the ODRA a contract dispute seeking an equitable 
adjustment or other damages after the contractor has accepted final 
contract payment, with the exception of FAA contract disputes related to 
warranty issues, gross mistakes amounting to fraud or latent defects. 
FAA contract disputes against the contractor based on warranty issues 
must be filed within the time specified under applicable contract 
warranty provisions. Any FAA contract disputes against the contractor 
based on gross mistakes amounting to fraud or latent defects shall be 
filed with the ODRA within two (2) years of the date on which the FAA 
knew or should have known of the presence of the fraud or latent defect.
    (d) A party shall serve a copy of the contract dispute upon the 
other party, by means reasonably calculated to be received on the same 
day as the filing is received by the ODRA.
    (e) With the exception of the time limitations established herein 
for the filing of contract disputes, the ODRA retains the discretion to 
modify any timeframe established herein in connection with contract 
disputes.



Sec. 17.29  Informal resolution period.

    (a) The ODRA process for contract disputes includes an informal 
resolution period of twenty (20) business days from the date of filing 
in order for the parties to attempt to informally resolve the contract 
dispute either through direct negotiation or with the assistance of the 
ODRA. The CO, with the advice of FAA legal counsel, has full discretion 
to settle contract disputes, except where the matter involves fraud.
    (b) During the informal resolution period, if the parties request 
it, the ODRA will appoint a DRO for ADR who will discuss ADR options 
with the parties, offer his or her services as a potential neutral, and 
assist the parties to enter into an agreement for a formal ADR process. 
A person serving as a neutral in an ADR effort in a matter shall not 
serve as an adjudicating DRO or Special Master for that matter.
    (c) The informal resolution period may be extended at the request of 
the parties for good cause.

[[Page 120]]

    (d) If the matter has not been resolved informally, the parties 
shall file joint or separate statements with the ODRA no later than 
twenty (20) business days after the filing of the contract dispute. The 
ODRA may extend this time, pursuant to Sec. 17.27(e). The statement(s) 
shall include either:
    (1) A joint request for ADR, or an executed ADR agreement, pursuant 
to Sec. 17.37(d), specifying which ADR techniques will be employed; or
    (2) Written explanation(s) as to why ADR proceedings will not be 
used and why the Adjudicative Process will be needed.
    (e) If the contract dispute is not completely resolved during the 
informal resolution period, the ODRA's Adjudicative Process will 
commence unless the parties have reached an agreement to attempt a 
formal ADR effort. As part of such an ADR agreement the parties, with 
the concurrence of the ODRA, may agree to defer commencement of the 
adjudication process pending completion of the ADR or that the ADR and 
adjudication process will run concurrently. If a formal ADR is attempted 
but does not completely resolve the contract dispute, the Adjudicative 
Process will commence.
    (f) The ODRA shall hold a status conference with the parties within 
ten (10) business days, or as soon thereafter as is practicable, of the 
ODRA's receipt of a written notification that ADR proceedings will not 
be used, or have not fully resolved the Contract Dispute. The purpose of 
the status conference will be to commence the Adjudicative Process and 
establish the schedule for adjudication.
    (g) The submission of a statement which indicates that ADR will not 
be utilized will not in any way preclude the parties from engaging in 
non-binding ADR techniques during the Adjudicative Process, pursuant to 
subpart D of this part.



Sec. 17.31  Dismissal or summary decision of contract disputes.

    (a) Any party may request by motion, or the ODRA on its own 
initiative may recommend or direct, that a contract dispute be 
dismissed, or that a count or portion thereof be stricken, if:
    (1) It was not timely filed;
    (2) It was filed by a subcontractor or other person or entity 
lacking standing;
    (3) It fails to state a matter upon which relief may be had; or
    (4) It involves a matter not subject to the jurisdiction of the 
ODRA.
    (b) Any party may request by motion, or the ODRA on its own 
initiative may recommend or direct, that a summary decision be issued 
with respect to a contract dispute, or any count or portion thereof if 
there are no material facts in dispute and a party is entitled to a 
summary decision as a matter of law.
    (c) In connection with any potential dismissal of a contract 
dispute, or summary decision, the ODRA will consider any material facts 
in dispute in a light most favorable to the party against whom the 
dismissal or summary decision would be entered, and draw all factual 
inferences in favor of that party.
    (d) At any time, whether pursuant to a motion or on its own 
initiative and at its discretion, the ODRA may:
    (1) Dismiss or strike a count or portion of a contract dispute or 
enter a partial summary decision;
    (2) Recommend to the Administrator that the entire contract dispute 
be dismissed or that a summary decision be entered; or
    (3) With a delegation from the Administrator, dismiss the entire 
contract dispute or enter a summary decision with respect to the entire 
contract dispute.
    (e) An order of dismissal of the entire contract dispute or summary 
decision with respect to the entire contract dispute, issued either by 
the Administrator or by the ODRA, on the grounds set forth in this 
section, shall constitute a final agency order. An ODRA order dismissing 
or striking a count or portion of a contract dispute or entering a 
partial summary judgment shall not constitute a final agency order, 
unless and until such ODRA order is incorporated or otherwise adopted in 
a final agency decision of the Administrator or the Administrator's 
delegee regarding the remainder of the dispute.
    (f) Prior to recommending or entering either a dismissal or a 
summary decision, either in whole or in part, the ODRA shall afford all 
parties against

[[Page 121]]

whom the dismissal or summary decision would be entered the opportunity 
to respond to a proposed dismissal or summary decision.



Sec. 17.33  Adjudicative Process for contract disputes.

    (a) The Adjudicative Process for contract disputes will be commenced 
by the ODRA Director upon being notified by the ADR neutral or by any 
party that either--
    (1) The parties will not be attempting ADR; or
    (2) The parties have not settled all of the dispute issues via ADR, 
and it is unlikely that they can do so within the time period allotted 
and/or any reasonable extension.
    (b) In cases initiated by a contractor against the FAA, within 
twenty (20) business days of the commencement of the Adjudicative 
Process or as scheduled by the ODRA, the Product Team shall prepare and 
submit to the ODRA, with a copy to the contractor, a chronologically 
arranged and indexed substantive response, containing a legal and 
factual position regarding the dispute and all documents relevant to the 
facts and issues in dispute. The contractor will be entitled, at a 
specified time, to supplement the record with additional documents.
    (c) In cases initiated by the FAA against a contractor, within 
twenty (20) business days of the commencement of the Adjudicative 
Process or as scheduled by the ODRA, the contractor shall prepare and 
submit to the ODRA, with a copy to the Product Team counsel, a 
chronologically arranged and indexed substantive response, containing a 
legal and factual position regarding the dispute and all documents 
relevant to the facts and issues in dispute. The Product Team will be 
entitled, at a specified time, to supplement the record with additional 
documents.
    (d) Unless timely objection is made, documents properly filed with 
the ODRA will be deemed admitted into the administrative record. 
Discovery requests and responses are not part of the record and will not 
be filed with the ODRA, except in connection with a motion or other 
permissible filing. Designated, relevant portions of such documents may 
be filed, with the permission of the ODRA.
    (e) The Director of the ODRA shall assign a DRO or a Special Master 
to conduct adjudicatory proceedings, develop the administrative 
adjudication record and prepare findings and recommendations for the 
review of the ODRA Director or the Director's designee.
    (f) The DRO or Special Master may conduct a status conference(s) as 
necessary and issue such orders or decisions as are necessary to promote 
the efficient resolution of the contract dispute.
    (g) At any such status conference, or as necessary during the 
Adjudicative Process, the DRO or Special Master will:
    (1) Determine the appropriate amount of discovery required;
    (2) Review the need for a protective order, and if one is needed, 
prepare a protective order pursuant to Sec. 17.9;
    (3) Determine whether any issue can be stricken; and
    (4) Prepare necessary procedural orders for the proceedings.
    (h) Unless otherwise provided by the DRO or Special Master, or by 
agreement of the parties with the concurrence of the DRO or Special 
Master, responses to written discovery shall be due within thirty (30) 
business days from the date received.
    (i) At a time or at times determined by the DRO or Special Master, 
and in advance of the decision of the case, the parties shall make 
individual final submissions to the ODRA and to the DRO or Special 
Master, which submissions shall include the following:
    (1) A statement of the issues;
    (2) A proposed statement of undisputed facts related to each issue 
together with citations to the administrative record or other supporting 
materials;
    (3) Separate statements of disputed facts related to each issue, 
with appropriate citations to documents in the Dispute File, to pages of 
transcripts of any hearing or deposition, or to any affidavit or exhibit 
which a party may wish to submit with its statement;
    (4) Separate legal analyses in support of the parties' respective 
positions on disputed issues.

[[Page 122]]

    (j) Each party shall serve a copy of its final submission on the 
other party by means reasonably calculated so that the other party 
receives such submissions on the same day it is received by the ODRA.
    (k) The DRO or Special Master may decide the contract dispute on the 
basis of the administrative record and the submissions referenced in 
this section, or may, in the DRO or Special Master's discretion, direct 
the parties to make additional presentations in writing. The DRO or 
Special Master may conduct hearings, and may limit the hearings to the 
testimony of specific witnesses and/or presentations regarding specific 
issues. The DRO or Special Master shall control the nature and conduct 
of all hearings, including the sequence and extent of any testimony. 
Evidentiary hearings on the record shall be conducted by the ODRA:
    (1) Where the DRO or Special Master determines that there are 
complex factual issues in dispute that cannot adequately or efficiently 
be developed solely by means of written presentations and/or that 
resolution of the controversy will be dependent on his/her assessment of 
the credibility of statements provided by individuals with first-hand 
knowledge of the facts; or
    (2) Upon request of any party to the contract dispute, unless the 
DRO or Special Master finds specifically that a hearing is unnecessary 
and that no party will be prejudiced by limiting the record in the 
adjudication to the parties' written submissions. All witnesses at any 
such hearing shall be subject to cross-examination by the opposing party 
and to questioning by the DRO or Special Master.
    (l) The DRO or Special Master shall prepare findings and 
recommendations, which will contain findings of fact, application of the 
principles of the AMS and other law or authority applicable to the 
findings of fact, and a recommendation for a final FAA order.
    (m) The DRO or Special Master shall conduct a de novo review using 
the preponderance of the evidence standard, unless a different standard 
is prescribed for a particular issue. Notwithstanding the above, 
allegations that government officials acted with bias or in bad faith 
must be established by clear and convincing evidence.
    (n) The Director of the ODRA may review the status of any contract 
dispute in the Adjudicative Process with the DRO or Special Master.
    (o) A DRO or Special Master shall submit findings and 
recommendations to the Director of the ODRA or the Director's designee. 
The findings and recommendations will be released to the parties and to 
the public, upon issuance of the final FAA order in the case. Should an 
ODRA protective order be issued in connection with the contract dispute, 
or should the matter involve proprietary or competition-sensitive 
information, a redacted version of the findings and recommendations 
omitting any protected information, shall be prepared wherever possible 
and released to the public, as soon as is practicable, along with a copy 
of the final FAA order. Only persons admitted by the ODRA under the 
protective order and Government personnel shall be provided copies of 
the unredacted findings and recommendations.
    (p) Attorneys' fees of a qualified prevailing contractor are 
allowable to the extent permitted by the EAJA, 5 U.S.C. 504(a)(1). See 
14 CFR part 14.
    (q) Other than communications regarding purely procedural matters or 
ADR, there shall be no substantive ex parte communication between ODRA 
personnel and any principal or representative of a party concerning a 
pending or potentially pending matter. A potential or serving ADR 
neutral may communicate on an ex parte basis to establish or conduct the 
ADR.



                Subpart D_Alternative Dispute Resolution



Sec. 17.35  Use of alternative dispute resolution.

    (a) By statutory mandate, it is the policy of the FAA to use 
voluntary ADR to the maximum extent practicable to resolve matters 
pending at the ODRA. The ODRA therefore uses voluntary ADR as its 
primary means of resolving all factual, legal, and procedural 
controversies.
    (b) The parties are encouraged to make a good faith effort to 
explore

[[Page 123]]

ADR possibilities in all cases and to employ ADR in every appropriate 
case. The ODRA uses ADR techniques such as mediation, neutral 
evaluation, binding arbitration or variations of these techniques as 
agreed by the parties and approved by the ODRA. At the beginning of each 
case, the ODRA assigns a DRO as a potential neutral to explore ADR 
options with the parties and to convene an ADR process. See Sec. 
17.35(b).
    (c) The ODRA Adjudicative Process will be used where the parties 
cannot achieve agreement on the use of ADR; where ADR has been employed 
but has not resolved all pending issues in dispute; or where the ODRA 
concludes that ADR will not provide an expeditious means of resolving a 
particular dispute. Even where the Adjudicative Process is to be used, 
the ODRA, with the parties' consent, may employ informal ADR techniques 
concurrently with the adjudication.



Sec. 17.37  Election of alternative dispute resolution process.

    (a) The ODRA will make its personnel available to serve as Neutrals 
in ADR proceedings and, upon request by the parties, will attempt to 
make qualified non-FAA personnel available to serve as Neutrals through 
neutral-sharing programs and other similar arrangements. The parties may 
elect to employ a mutually acceptable compensated neutral at their 
expense.
    (b) The parties using an ADR process to resolve a protest shall 
submit an executed ADR agreement containing the information outlined in 
paragraph (d) of this section to the ODRA pursuant to Sec. 17.17(c). 
The ODRA may extend this time for good cause.
    (c) The parties using an ADR process to resolve a contract dispute 
shall submit an executed ADR agreement containing the information 
outlined in paragraph (d) of this section to the ODRA pursuant to Sec. 
17.29.
    (d) The parties to a protest or contract dispute who elect to use 
ADR must submit to the ODRA an ADR agreement setting forth:
    (1) The agreed ADR procedures to be used; and
    (2) The name of the neutral. If a compensated neutral is to be used, 
the agreement must address how the cost of the neutral's services will 
be reimbursed.
    (e) Non-binding ADR techniques are not mutually exclusive, and may 
be used in combination if the parties agree that a combination is most 
appropriate to the dispute. The techniques to be employed must be 
determined in advance by the parties and shall be expressly described in 
their ADR agreement. The agreement may provide for the use of any fair 
and reasonable ADR technique that is designed to achieve a prompt 
resolution of the matter. An ADR agreement for non-binding ADR shall 
provide for a termination of ADR proceedings and the commencement of 
adjudication under the Adjudicative Process, upon the election of any 
party. Notwithstanding such termination, the parties may still engage 
with the ODRA in ADR techniques (neutral evaluation and/or informal 
mediation) concurrently with adjudication.
    (f) Binding arbitration is available through the ODRA, subject to 
the provisions of applicable law and the ODRA Binding Arbitration 
Guidance dated October 2001 as developed in consultation with the 
Department of Justice.
    (g) The parties may, where appropriate in a given case, submit to 
the ODRA a negotiated protective order for use in ADR in accordance with 
the requirements of Sec. 17.9.



Sec. 17.39  Confidentiality of ADR.

    (a) The provisions of the Administrative Dispute Resolution Act of 
1996, 5 U.S.C. 571, et seq., shall apply to ODRA ADR proceedings.
    (b) The ODRA looks to the principles of Rule 408 of the Federal 
Rules of Evidence in deciding admissibility issues related to ADR 
communications.
    (c) ADR communications are not part of the administrative record 
unless otherwise agreed by the parties.



                      Subpart E_Finality and Review



Sec. 17.41  Final orders.

    All final FAA orders regarding protests or contract disputes under 
this part are to be issued by the FAA Administrator or by a delegee of 
the Administrator.

[[Page 124]]



Sec. 17.43  Judicial review.

    (a) A protester or contractor may seek review of a final FAA order, 
pursuant to 49 U.S.C. 46110, only after the administrative remedies of 
this part have been exhausted.
    (b) A copy of the petition for review shall be filed with the ODRA 
and the FAA Chief Counsel on the date that the petition for review is 
filed with the appropriate circuit court of appeals.



Sec. 17.45  Conforming amendments.

    The FAA shall amend pertinent provisions of the AMS, standard 
contract forms and clauses, and any guidance to contracting officials, 
so as to conform to the provisions of this part.



Sec. 17.47  Reconsideration.

    The ODRA will not entertain requests for reconsideration as a 
routine matter, or where such requests evidence mere disagreement with a 
decision or restatements of previous arguments. A party seeking 
reconsideration must demonstrate either clear errors of fact or law in 
the underlying decision or previously unavailable evidence that warrants 
reversal or modification of the decision. In order to be considered, 
requests for reconsideration must be filed within ten (10) business days 
of the date of issuance of the public version of the subject decision or 
order.



                         Subpart F_Other Matters



Sec. 17.49  Sanctions.

    If any party or its representative fails to comply with an Order or 
Directive of the ODRA, the ODRA may enter such orders and take such 
other actions as it deems necessary and in the interest of justice.



Sec. 17.51  Decorum and professional conduct.

    Legal representatives are expected to conduct themselves at all 
times in a civil and respectful manner appropriate to an administrative 
forum. Additionally, counsel are expected to conduct themselves at all 
times in a professional manner and in accordance with all applicable 
rules of professional conduct.



Sec. 17.53  Orders and subpoenas for testimony and document production.

    (a) Parties are encouraged to seek cooperative and voluntary 
production of documents and witnesses prior to requesting a subpoena or 
an order under this section.
    (b) Upon request by a party, or on his or her own initiative, a DRO 
or Special Master may, for good cause shown, order a person to give 
testimony by deposition and to produce records. Section 46104(c) of 
Title 49 of the United States Code governs the conduct of depositions or 
document production.
    (c) Upon request by a party, or on his or her own initiative, a DRO 
or Special Master may, for good cause shown, subpoena witnesses or 
records related to a hearing from any place in the United States to the 
designated place of a hearing.
    (d) A subpoena or order under this section may be served by a United 
States marshal or deputy marshal, or by any other person who is not a 
party and not less than 18 years of age. Service upon a person named 
therein shall be made by personally delivering a copy to that person and 
tendering the fees for one day's attendance and the mileage provided by 
28 U.S.C. 1821 or other applicable law; however, where the subpoena is 
issued on behalf of the Product Team, money payments need not be 
tendered in advance of attendance. The person serving the subpoena or 
order shall file a declaration of service with the ODRA, executed in the 
form required by 28 U.S.C. 1746. The declaration of service shall be 
filed promptly with the ODRA, and before the date on which the person 
served must respond to the subpoena or order.
    (e) Upon written motion by the person subpoenaed or ordered under 
this section, or by a party, made within ten (10) business days after 
service, but in any event not later than the time specified in the 
subpoena or order for compliance, the DRO may--
    (1) Rescind or modify the subpoena or order if it is unreasonable 
and oppressive or for other good cause shown, or
    (2) Require the party on whose behalf the subpoena or order was 
issued to advance the reasonable cost of producing documentary evidence. 
Where circumstances require, the DRO may act

[[Page 125]]

upon such a motion at any time after a copy has been served upon all 
parties.
    (f) The party that requests the DRO to issue a subpoena or order 
under this section shall be responsible for the payment of fees and 
mileage, as required by 49 U.S.C. 46104(d), for witnesses, officers who 
serve the order, and the officer before whom a deposition is taken.
    (g) Subpoenas and orders issued under this section may be enforced 
in a judicial proceeding under 49 U.S.C. 46104(b).



Sec. 17.55  Standing orders of the ODRA Director.

    The Director may issue such Standing Orders as necessary for the 
orderly conduct of business before the ODRA.



                         Subpart G_Pre-Disputes



Sec. 17.57  Dispute resolution process for Pre-disputes.

    (a) All potential disputes arising under contracts or solicitations 
with the FAA may be resolved with the consent of the parties to the 
dispute under this subpart.
    (b) Pre-disputes shall be filed with the ODRA pursuant to Sec. 
17.59.
    (c) The time limitations for the filing of Protests and Contract 
Disputes established in Sec. Sec. 17.15(a) and 17.27(c) will not be 
extended by efforts to resolve the dispute under this subpart.



Sec. 17.59  Filing a Pre-dispute.

    (a) A Pre-dispute must be in writing, affirmatively state that it is 
a Pre-dispute pursuant to this subpart, and shall contain:
    (1) The party's name, address, telephone and Fax numbers and the 
name, address, telephone and Fax numbers of the contractor's legal 
representative(s) (if any);
    (2) The contract or solicitation number and the name of the 
Contracting Officer;
    (3) A chronological statement of the facts and of the legal grounds 
for the party's positions regarding the dispute citing to relevant 
contract or solicitation provisions and documents and attaching copies 
of those provisions and documents; and
    (4) The signature of a duly authorized legal representative of the 
initiating party.
    (b) Pre-disputes shall be filed at the following address: ODRA, AGC-
70, Federal Aviation Administration, 800 Independence Avenue, SW., Room 
323, Washington, DC 20591; Telephone: (202) 267-3290, Fax: (202) 267-
3720.
    (c) Upon the filing of a Pre-dispute with the ODRA, the ODRA will 
contact the opposing party to offer its services pursuant to Sec. 
17.57. If the opposing party agrees, the ODRA will provide Pre-dispute 
services. If the opposing party does not agree, the ODRA Pre-dispute 
file will be closed and no service will be provided.



Sec. 17.61  Use of alternative dispute resolution.

    (a) Only non-binding, voluntary ADR will be used to attempt to 
resolve a Pre-dispute pursuant to Sec. 17.37.
    (b) ADR conducted under this subpart is subject to the 
confidentiality requirements of Sec. 17.39.



    Sec. Appendix A to Part 17--Alternative Dispute Resolution (ADR)

    A. The FAA dispute resolution procedures encourage the parties to 
protests and contract disputes to use ADR as the primary means to 
resolve protests and contract disputes, pursuant to the Administrative 
Dispute Resolution Act of 1996, Public Law 104-320, 5 U.S.C. 570-579, 
and Department of Transportation and FAA policies to utilize ADR to the 
maximum extent practicable. Under the procedures presented in this part, 
the ODRA encourages parties to consider ADR techniques such as case 
evaluation, mediation, or arbitration.
    B. ADR encompasses a number of processes and techniques for 
resolving protests or contract disputes. The most commonly used types 
include:
    (1) Mediation. The neutral or compensated neutral ascertains the 
needs and interests of both parties and facilitates discussions between 
or among the parties and an amicable resolution of their differences, 
seeking approaches to bridge the gaps between the parties'' respective 
positions. The neutral or compensated neutral can meet with the parties 
separately, conduct joint meetings with the parties'' representatives, 
or employ both methods in appropriate cases.
    (2) Neutral Evaluation. At any stage during the ADR process, as the 
parties may agree, the neutral or compensated neutral will provide a 
candid assessment and opinion of the

[[Page 126]]

strengths and weaknesses of the parties'' positions as to the facts and 
law, so as to facilitate further discussion and resolution.
    (3) Binding Arbitration. The ODRA, after consultation with the 
United States Department of Justice in accordance with the provisions of 
the Administrative Disputes Resolution Act offers true binding 
arbitration in cases within its jurisdiction. The ODRA's Guidance for 
the Use of Binding Arbitration may be found on its website at: http://
www.faa.gov/go/odra.

[[Page 127]]



                          SUBCHAPTER C_AIRCRAFT





PART 21_CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS--
Table of Contents



Special Federal Aviation Regulation No. 88

                            Subpart A_General

Sec.
21.1 Applicability and definitions.
21.2 Falsification of applications, reports, or records.
21.3 Reporting of failures, malfunctions, and defects.
21.4 ETOPS reporting requirements.
21.5 Airplane or Rotorcraft Flight Manual.
21.6 Manufacture of new aircraft, aircraft engines, and propellers.
21.7 Continued airworthiness and safety improvements for transport 
          category airplanes.
21.8 Approval of articles.
21.9 Replacement and modification articles.

                       Subpart B_Type Certificates

21.11 Applicability.
21.13 Eligibility.
21.15 Application for type certificate.
21.16 Special conditions.
21.17 Designation of applicable regulations.
21.19 Changes requiring a new type certificate.
21.20 Compliance with applicable requirements.
21.21 Issue of type certificate: normal, utility, acrobatic, commuter, 
          and transport category aircraft; manned free balloons; special 
          classes of aircraft; aircraft engines; propellers.
21.23 [Reserved]
21.24 Issuance of type certificate: primary category aircraft.
21.25 Issue of type certificate: Restricted category aircraft.
21.27 Issue of type certificate: surplus aircraft of the Armed Forces.
21.29 Issue of type certificate: import products.
21.31 Type design.
21.33 Inspection and tests.
21.35 Flight tests.
21.37 Flight test pilot.
21.39 Flight test instrument calibration and correction report.
21.41 Type certificate.
21.43 Location of manufacturing facilities.
21.45 Privileges.
21.47 Transferability.
21.49 Availability.
21.50 Instructions for continued airworthiness and manufacturer's 
          maintenance manuals having airworthiness limitations sections.
21.51 Duration.
21.53 Statement of conformity.
21.55 Responsibility of type certificate holders to provide written 
          licensing agreements.

                 Subpart C_Provisional Type Certificates

21.71 Applicability.
21.73 Eligibility.
21.75 Application.
21.77 Duration.
21.79 Transferability.
21.81 Requirements for issue and amendment of Class I provisional type 
          certificates.
21.83 Requirements for issue and amendment of Class II provisional type 
          certificates.
21.85 Provisional amendments to type certificates.

                 Subpart D_Changes to Type Certificates

21.91 Applicability.
21.93 Classification of changes in type design.
21.95 Approval of minor changes in type design.
21.97 Approval of major changes in type design.
21.99 Required design changes.
21.101 Designation of applicable regulations.

                Subpart E_Supplemental Type Certificates

21.111 Applicability.
21.113 Requirement for supplemental type certificate.
21.115 Applicable requirements.
21.117 Issue of supplemental type certificates.
21.119 Privileges.
21.120 Responsibility of supplemental type certificate holders to 
          provide written permission for alterations.

               Subpart F_Production Under Type Certificate

21.121 Applicability.
21.122 Location of or change to manufacturing facilities.
21.123 Production under type certificate.
21.125 [Reserved]
21.127 Tests: aircraft.
21.128 Tests: aircraft engines.
21.129 Tests: propellers.
21.130 Statement of conformity.

[[Page 128]]

                    Subpart G_Production Certificates

21.131 Applicability.
21.132 Eligibility.
21.133 Application.
21.135 Organization.
21.137 Quality system.
21.138 Quality manual.
21.139 Location of or change to manufacturing facilities.
21.140 Inspections and tests.
21.141 Issuance.
21.142 Production limitation record.
21.143 Duration.
21.144 Transferability.
21.145 Privileges.
21.146 Responsibility of holder.
21.147 Amendment of production certificates.
21.150 Changes in quality system.

                  Subpart H_Airworthiness Certificates

21.171 Applicability.
21.173 Eligibility.
21.175 Airworthiness certificates: classification.
21.177 Amendment or modification.
21.179 Transferability.
21.181 Duration.
21.182 Aircraft identification.
21.183 Issue of standard airworthiness certificates for normal, utility, 
          acrobatic, commuter, and transport category aircraft; manned 
          free balloons; and special classes of aircraft.
21.184 Issue of special airworthiness certificates for primary category 
          aircraft.
21.185 Issue of airworthiness certificates for restricted category 
          aircraft.
21.187 Issue of multiple airworthiness certification.
21.189 Issue of airworthiness certificate for limited category aircraft.
21.190 Issue of a special airworthiness certificate for a light-sport 
          category aircraft.
21.191 Experimental certificates.
21.193 Experimental certificates: general.
21.195 Experimental certificates: Aircraft to be used for market 
          surveys, sales demonstrations, and customer crew training.
21.197 Special flight permits.
21.199 Issue of special flight permits.

            Subpart I_Provisional Airworthiness Certificates

21.211 Applicability.
21.213 Eligibility.
21.215 Application.
21.217 Duration.
21.219 Transferability.
21.221 Class I provisional airworthiness certificates.
21.223 Class II provisional airworthiness certificates.
21.225 Provisional airworthiness certificates corresponding with 
          provisional amendments to type certificates.

Subpart J [Reserved]

                 Subpart K_Parts Manufacturer Approvals

21.301 Applicability.
21.303 Application.
21.305 Organization.
21.307 Quality system.
21.308 Quality manual.
21.309 Location of or change to manufacturing facilities.
21.310 Inspections and tests.
21.311 Issuance.
21.313 Duration.
21.314 Transferability.
21.316 Responsibility of holder.
21.319 Design changes.
21.320 Changes in quality system.

                Subpart L_Export Airworthiness Approvals

21.321 Applicability.
21.325 Export airworthiness approvals.
21.327 Application.
21.329 Issuance of export certificates of airworthiness.
21.331 Issuance of export airworthiness approvals for aircraft engines, 
          propellers, and articles.
21.335 Responsibilities of exporters.

Subpart M [Reserved]

 Subpart N_Acceptance of Aircraft Engines, Propellers, and Articles for 
                                 Import

21.500 Acceptance of aircraft engines and propellers.
21.502 Acceptance of articles.

              Subpart O_Technical Standard Order Approvals

21.601 Applicability and definitions.
21.603 Application.
21.605 Organization.
21.607 Quality system.
21.608 Quality manual.
21.609 Location of or change to manufacturing facilities.
21.610 Inspections and tests.
21.611 Issuance.
21.613 Duration.
21.614 Transferability.
21.616 Responsibility of holder.
21.618 Approval for deviation.
21.619 Design changes.
21.620 Changes in quality system.
21.621 Issue of letters of TSO design approval: Import articles.

[[Page 129]]

             Subpart P_Special Federal Aviation Regulations

21.700 SFAR No. 111--Lavatory oxygen systems.

    Authority: 42 U.S.C. 7572; 49 U.S.C. 106(g), 40105, 40113, 44701-
44702, 44704, 44707, 44709, 44711, 44713, 44715, 45303.

    Editorial Note: For miscellaneous amendments to cross references in 
this Part 21 see Amdt. 21-10, 31 FR 9211, July 6, 1966.

    Editorial Note: Nomenclature changes to part 21 appear at 74 FR 
53384, Oct. 16, 2009.



Sec. Special Federal Aviation Regulation No. 88--Fuel Tank System Fault 
                    Tolerance Evaluation Requirements

    1. Applicability. This SFAR applies to the holders of type 
certificates, and supplemental type certificates that may affect the 
airplane fuel tank system, for turbine-powered transport category 
airplanes, provided the type certificate was issued after January 1, 
1958, and the airplane has either a maximum type certificated passenger 
capacity of 30 or more, or a maximum type certificated payload capacity 
of 7,500 pounds or more. This SFAR also applies to applicants for type 
certificates, amendments to a type certificate, and supplemental type 
certificates affecting the fuel tank systems for those airplanes 
identified above, if the application was filed before June 6, 2001, the 
effective date of this SFAR, and the certificate was not issued before 
June 6, 2001.
    2. Compliance: Each type certificate holder, and each supplemental 
type certificate holder of a modification affecting the airplane fuel 
tank system, must accomplish the following within the compliance times 
specified in paragraph (e) of this section:
    (a) Conduct a safety review of the airplane fuel tank system to 
determine that the design meets the requirements of Sec. Sec. 25.901 
and 25.981(a) and (b) of this chapter. If the current design does not 
meet these requirements, develop all design changes to the fuel tank 
system that are necessary to meet these requirements. The FAA (Aircraft 
Certification Office (ACO), or office of the Transport Airplane 
Directorate, having cognizance over the type certificate for the 
affected airplane) may grant an extension of the 18-month compliance 
time for development of design changes if:
    (1) The safety review is completed within the compliance time;
    (2) Necessary design changes are identified within the compliance 
time; and
    (3) Additional time can be justified, based on the holder's 
demonstrated aggressiveness in performing the safety review, the 
complexity of the necessary design changes, the availability of interim 
actions to provide an acceptable level of safety, and the resulting 
level of safety.
    (b) Develop all maintenance and inspection instructions necessary to 
maintain the design features required to preclude the existence or 
development of an ignition source within the fuel tank system of the 
airplane.
    (c) Submit a report for approval to the FAA Aircraft Certification 
Office (ACO), or office of the Transport Airplane Directorate, having 
cognizance over the type certificate for the affected airplane, that:
    (1) Provides substantiation that the airplane fuel tank system 
design, including all necessary design changes, meets the requirements 
of Sec. Sec. 25.901 and 25.981(a) and (b) of this chapter; and
    (2) Contains all maintenance and inspection instructions necessary 
to maintain the design features required to preclude the existence or 
development of an ignition source within the fuel tank system throughout 
the operational life of the airplane.
    (d) The Aircraft Certification Office (ACO), or office of the 
Transport Airplane Directorate, having cognizance over the type 
certificate for the affected airplane, may approve a report submitted in 
accordance with paragraph 2(c) if it determines that any provisions of 
this SFAR not complied with are compensated for by factors that provide 
an equivalent level of safety.
    (e) Each type certificate holder must comply no later than December 
6, 2002, or within 18 months after the issuance of a type certificate 
for which application was filed before June 6, 2001, whichever is later; 
and each supplemental type certificate holder of a modification 
affecting the airplane fuel tank system must comply no later than June 
6, 2003, or within 18 months after the issuance of a supplemental type 
certificate for which application was filed before June 6, 2001, 
whichever is later.

[Doc. No. 1999-6411, 66 FR 23129, May 7, 2001, as amended by Amdt. 21-
82, 67 FR 57493, Sept. 10, 2002; 67 FR 70809, Nov. 26, 2002; Amdt. 21-
83, 67 FR 72833, Dec. 9, 2002]



                            Subpart A_General



Sec. 21.1  Applicability and definitions.

    (a) This part prescribes--
    (1) Procedural requirements for issuing and changing--
    (i) Design approvals;
    (ii) Production approvals;
    (iii) Airworthiness certificates; and
    (iv) Airworthiness approvals;
    (2) Rules governing applicants for, and holders of, any approval or 
certificate specified in paragraph (a)(1) of this section; and

[[Page 130]]

    (3) Procedural requirements for the approval of articles.
    (b) For the purposes of this part--
    (1) Airworthiness approval means a document issued by the FAA for an 
aircraft, aircraft engine, propeller, or article which certifies that 
the aircraft, aircraft engine, propeller, or article conforms to its 
approved design and is in a condition for safe operation;
    (2) Article means a material, part, component, process, or 
appliance;
    (3) Commercial part means an article that is listed on an FAA-
approved Commercial Parts List included in a design approval holder's 
Instructions for Continued Airworthiness required by Sec. 21.50;
    (4) Design approval means a type certificate (including amended and 
supplemental type certificates) or the approved design under a PMA, TSO 
authorization, letter of TSO design approval, or other approved design;
    (5) Product means an aircraft, aircraft engine, or propeller;
    (6) Production approval means a document issued by the FAA to a 
person that allows the production of a product or article in accordance 
with its approved design and approved quality system, and can take the 
form of a production certificate, a PMA, or a TSO authorization;
    (7) State of Design means the country or jurisdiction having 
regulatory authority over the organization responsible for the design 
and continued airworthiness of a civil aeronautical product or article;
    (8) State of Manufacture means the country or jurisdiction having 
regulatory authority over the organization responsible for the 
production and airworthiness of a civil aeronautical product or article.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53384, Oct. 16, 2009]



Sec. 21.2  Falsification of applications, reports, or records.

    (a) A person may not make or cause to be made--
    (1) Any fraudulent, intentionally false, or misleading statement on 
any application for a certificate or approval under this part;
    (2) Any fraudulent, intentionally false, or misleading statement in 
any record or report that is kept, made, or used to show compliance with 
any requirement of this part;
    (3) Any reproduction for a fraudulent purpose of any certificate or 
approval issued under this part.
    (4) Any alteration of any certificate or approval issued under this 
part.
    (b) The commission by any person of an act prohibited under 
paragraph (a) of this section is a basis for--
    (1) Denying issuance of any certificate or approval under this part; 
and
    (2) Suspending or revoking any certificate or approval issued under 
this part and held by that person.

[Doc. No. 23345, 57 FR 41367, Sept. 9, 1992, as amended by Amdt. 21-92, 
74 FR 53384, Oct. 16, 2009; Amdt. 21-92A, 75FR 9095, Mar. 1, 2010]



Sec. 21.3  Reporting of failures, malfunctions, and defects.

    (a) The holder of a type certificate (including amended or 
supplemental type certificates), a PMA, or a TSO authorization, or the 
licensee of a type certificate must report any failure, malfunction, or 
defect in any product or article manufactured by it that it determines 
has resulted in any of the occurrences listed in paragraph (c) of this 
section.
    (b) The holder of a type certificate (including amended or 
supplemental type certificates), a PMA, or a TSO authorization, or the 
licensee of a type certificate must report any defect in any product or 
article manufactured by it that has left its quality system and that it 
determines could result in any of the occurrences listed in paragraph 
(c) of this section.
    (c) The following occurrences must be reported as provided in 
paragraphs (a) and (b) of this section:
    (1) Fires caused by a system or equipment failure, malfunction, or 
defect.
    (2) An engine exhaust system failure, malfunction, or defect which 
causes damage to the engine, adjacent aircraft structure, equipment, or 
components.
    (3) The accumulation or circulation of toxic or noxious gases in the 
crew compartment or passenger cabin.
    (4) A malfunction, failure, or defect of a propeller control system.
    (5) A propeller or rotorcraft hub or blade structural failure.

[[Page 131]]

    (6) Flammable fluid leakage in areas where an ignition source 
normally exists.
    (7) A brake system failure caused by structural or material failure 
during operation.
    (8) A significant aircraft primary structural defect or failure 
caused by any autogenous condition (fatigue, understrength, corrosion, 
etc.).
    (9) Any abnormal vibration or buffeting caused by a structural or 
system malfunction, defect, or failure.
    (10) An engine failure.
    (11) Any structural or flight control system malfunction, defect, or 
failure which causes an interference with normal control of the aircraft 
for which derogates the flying qualities.
    (12) A complete loss of more than one electrical power generating 
system or hydraulic power system during a given operation of the 
aircraft.
    (13) A failure or malfunction of more than one attitude, airspeed, 
or altitude instrument during a given operation of the aircraft.
    (d) The requirements of paragraph (a) of this section do not apply 
to--
    (1) Failures, malfunctions, or defects that the holder of a type 
certificate (including amended or supplemental type certificates), PMA, 
TSO authorization, or the licensee of a type certificate determines--
    (i) Were caused by improper maintenance or use;
    (ii) Were reported to the FAA by another person under this chapter; 
or
    (iii) Were reported under the accident reporting provisions of 49 
CFR part 830 of the regulations of the National Transportation Safety 
Board.
    (2) Failures, malfunctions, or defects in products or articles--
    (i) Manufactured by a foreign manufacturer under a U.S. type 
certificate issued under Sec. 21.29 or under an approval issued under 
Sec. 21.621; or
    (ii) Exported to the United States under Sec. 21.502.
    (e) Each report required by this section--
    (1) Must be made to the Aircraft Certification Office in the region 
in which the person required to make the report is located within 24 
hours after it has determined that the failure, malfunction, or defect 
required to be reported has occurred. However, a report that is due on a 
Saturday or a Sunday may be delivered on the following Monday and one 
that is due on a holiday may be delivered on the next workday;
    (2) Must be transmitted in a manner and form acceptable to the FAA 
and by the most expeditious method available; and
    (3) Must include as much of the following information as is 
available and applicable:
    (i) The applicable product and article identification information 
required by part 45 of this chapter;
    (ii) Identification of the system involved; and
    (iii) Nature of the failure, malfunction, or defect.
    (f) If an accident investigation or service difficulty report shows 
that a product or article manufactured under this part is unsafe because 
of a manufacturing or design data defect, the holder of the production 
approval for that product or article must, upon request of the FAA, 
report to the FAA the results of its investigation and any action taken 
or proposed by the holder of that production approval to correct that 
defect. If action is required to correct the defect in an existing 
product or article, the holder of that production approval must send the 
data necessary for issuing an appropriate airworthiness directive to the 
appropriate aircraft certification office.

[Amdt. 21-36, 35 FR 18187, Nov. 28, 1970, as amended by Amdt. 21-37, 35 
FR 18450, Dec. 4, 1970; Amdt. 21-50, 45 FR 38346, June 9, 1980; Amdt. 
21-67, 54 FR 39291, Sept. 25, 1989; Amdt. 21-92, 74 FR 53385, Oct. 16, 
2009]



Sec. 21.4  ETOPS reporting requirements.

    (a) Early ETOPS: reporting, tracking, and resolving problems. The 
holder of a type certificate for an airplane-engine combination approved 
using the Early ETOPS method specified in part 25, Appendix K, of this 
chapter must use a system for reporting, tracking, and resolving each 
problem resulting in one of the occurrences specified in paragraph 
(a)(6) of this section.
    (1) The system must identify how the type certificate holder will 
promptly identify problems, report them to the responsible FAA aircraft 
certification office, and propose a solution to the

[[Page 132]]

FAA to resolve each problem. A proposed solution must consist of--
    (i) A change in the airplane or engine type design;
    (ii) A change in a manufacturing process;
    (iii) A change in an operating or maintenance procedure; or
    (iv) Any other solution acceptable to the FAA.
    (2) For an airplane with more than two engines, the system must be 
in place for the first 250,000 world fleet engine-hours for the approved 
airplane-engine combination.
    (3) For two-engine airplanes, the system must be in place for the 
first 250,000 world fleet engine-hours for the approved airplane-engine 
combination and after that until--
    (i) The world fleet 12-month rolling average IFSD rate is at or 
below the rate required by paragraph (b)(2) of this section; and
    (ii) The FAA determines that the rate is stable.
    (4) For an airplane-engine combination that is a derivative of an 
airplane-engine combination previously approved for ETOPS, the system 
need only address those problems specified in the following table, 
provided the type certificate holder obtains prior authorization from 
the FAA:

------------------------------------------------------------------------
                                          Then the Problem Tracking and
  If the change does not require a new    Resolution System must address
  airplane type certificate and . . .                 . . .
------------------------------------------------------------------------
(i) Requires a new engine type           All problems applicable to the
 certificate.                             new engine installation, and
                                          for the remainder of the
                                          airplane, problems in changed
                                          systems only.
(ii) Does not require a new engine type  Problems in changed systems
 certificate.                             only.
------------------------------------------------------------------------

    (5) The type certificate holder must identify the sources and 
content of data that it will use for its system. The data must be 
adequate to evaluate the specific cause of any in-service problem 
reportable under this section or Sec. 21.3(c) that could affect the 
safety of ETOPS.
    (6) In implementing this system, the type certificate holder must 
report the following occurrences:
    (i) IFSDs, except planned IFSDs performed for flight training.
    (ii) For two-engine airplanes, IFSD rates.
    (iii) Inability to control an engine or obtain desired thrust or 
power.
    (iv) Precautionary thrust or power reductions.
    (v) Degraded ability to start an engine in flight.
    (vi) Inadvertent fuel loss or unavailability, or uncorrectable fuel 
imbalance in flight.
    (vii) Turn backs or diversions for failures, malfunctions, or 
defects associated with an ETOPS group 1 significant system.
    (viii) Loss of any power source for an ETOPS group 1 significant 
system, including any power source designed to provide backup power for 
that system.
    (ix) Any event that would jeopardize the safe flight and landing of 
the airplane on an ETOPS flight.
    (x) Any unscheduled engine removal for a condition that could result 
in one of the reportable occurrences listed in this paragraph.
    (b) Reliability of two-engine airplanes--(1) Reporting of two-engine 
airplane in-service reliability. The holder of a type certificate for an 
airplane approved for ETOPS and the holder of a type certificate for an 
engine installed on an airplane approved for ETOPS must report monthly 
to their respective FAA type certificate holding office on the 
reliability of the world fleet of those airplanes and engines. The 
report provided by both the airplane and engine type certificate holders 
must address each airplane-engine combination approved for ETOPS. The 
FAA may approve quarterly reporting if the airplane-engine combination 
demonstrates an IFSD rate at or below those specified in paragraph 
(b)(2) of this section for a period acceptable to the FAA. This 
reporting may be combined with the reporting required by Sec. 21.3. The 
responsible type certificate holder must investigate any cause of an 
IFSD resulting from an occurrence

[[Page 133]]

attributable to the design of its product and report the results of that 
investigation to its FAA office responsible for administering its type 
certificate. Reporting must include:
    (i) Engine IFSDs, except planned IFSDs performed for flight 
training.
    (ii) The world fleet 12-month rolling average IFSD rates for all 
causes, except planned IFSDs performed for flight training.
    (iii) ETOPS fleet utilization, including a list of operators, their 
ETOPS diversion time authority, flight hours, and cycles.
    (2) World fleet IFSD rate for two-engine airplanes. The holder of a 
type certificate for an airplane approved for ETOPS and the holder of a 
type certificate for an engine installed on an airplane approved for 
ETOPS must issue service information to the operators of those airplanes 
and engines, as appropriate, to maintain the world fleet 12-month 
rolling average IFSD rate at or below the following levels:
    (i) A rate of 0.05 per 1,000 world-fleet engine-hours for an 
airplane-engine combination approved for up to and including 120-minute 
ETOPS. When all ETOPS operators have complied with the corrective 
actions required in the configuration, maintenance and procedures (CMP) 
document as a condition for ETOPS approval, the rate to be maintained is 
at or below 0.02 per 1,000 world-fleet engine-hours.
    (ii) A rate of 0.02 per 1,000 world-fleet engine-hours for an 
airplane-engine combination approved for up to and including 180-minute 
ETOPS, including airplane-engine combinations approved for 207-minute 
ETOPS in the North Pacific operating area under appendix P, section I, 
paragraph (h), of part 121 of this chapter.
    (iii) A rate of 0.01 per 1,000 world-fleet engine-hours for an 
airplane-engine combination approved for ETOPS beyond 180 minutes, 
excluding airplane-engine combinations approved for 207-minute ETOPS in 
the North Pacific operating area under appendix P, section I, paragraph 
(h), of part 121 of this chapter.

[Doc. No. FAA-2002-6717, 72 FR 1872, Jan. 16, 2007]



Sec. 21.5  Airplane or Rotorcraft Flight Manual.

    (a) With each airplane or rotorcraft not type certificated with an 
Airplane or Rotorcraft Flight Manual and having no flight time before 
March 1, 1979, the holder of a type certificate (including amended or 
supplemental type certificates) or the licensee of a type certificate 
must make available to the owner at the time of delivery of the aircraft 
a current approved Airplane or Rotorcraft Flight Manual.
    (b) The Airplane or Rotorcraft Flight Manual required by paragraph 
(a) of this section must contain the following information:
    (1) The operating limitations and information required to be 
furnished in an Airplane or Rotorcraft Flight Manual or in manual 
material, markings, and placards, by the applicable regulations under 
which the airplane or rotorcraft was type certificated.
    (2) The maximum ambient atmospheric temperature for which engine 
cooling was demonstrated must be stated in the performance information 
section of the Flight Manual, if the applicable regulations under which 
the aircraft was type certificated do not require ambient temperature on 
engine cooling operating limitations in the Flight Manual.

[Amdt. 21-46, 43 FR 2316, Jan. 16, 1978, as amended by Amdt. 21-92, 74 
FR 53385, Oct. 16, 2009]



Sec. 21.6  Manufacture of new aircraft, aircraft engines, and propellers.

    (a) Except as specified in paragraphs (b) and (c) of this section, 
no person may manufacture a new aircraft, aircraft engine, or propeller 
based on a type certificate unless the person--
    (1) Is the holder of the type certificate or has a licensing 
agreement from the holder of the type certificate to manufacture the 
product; and
    (2) Meets the requirements of subpart F or G of this part.
    (b) A person may manufacture one new aircraft based on a type 
certificate without meeting the requirements of paragraph (a) of this 
section if that person can provide evidence acceptable to the FAA that 
the manufacture of the aircraft by that person began before August 5, 
2004.

[[Page 134]]

    (c) The requirements of this section do not apply to--
    (1) New aircraft imported under the provisions of Sec. Sec. 
21.183(c), 21.184(b), or 21.185(c); and
    (2) New aircraft engines or propellers imported under the provisions 
of Sec. 21.500.

[Doc. No. FAA-2003-14825, 71 FR 52258, Sept. 1, 2006]



Sec. 21.7  Continued airworthiness and safety improvements for
transport category airplanes.

    (a) On or after December 10, 2007, the holder of a design approval 
and an applicant for a design approval must comply with the applicable 
continued airworthiness and safety improvement requirements of part 26 
of this subchapter.
    (b) For new transport category airplanes manufactured under the 
authority of the FAA, the holder or licensee of a type certificate must 
meet the applicable continued airworthiness and safety improvement 
requirements specified in part 26 of this subchapter for new production 
airplanes. Those requirements only apply if the FAA has jurisdiction 
over the organization responsible for final assembly of the airplane.

[Doc. No. FAA-2004-18379, Amdt. 21-90, 72 FR 63404, Nov. 8, 2007]



Sec. 21.8  Approval of articles.

    If an article is required to be approved under this chapter, it may 
be approved--
    (a) Under a PMA;
    (b) Under a TSO;
    (c) In conjunction with type certification procedures for a product; 
or
    (d) In any other manner approved by the FAA.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53385, Oct. 16, 2009]



Sec. 21.9  Replacement and modification articles.

    (a) If a person knows, or should know, that a replacement or 
modification article is reasonably likely to be installed on a type-
certificated product, the person may not produce that article unless it 
is--
    (1) Produced under a type certificate;
    (2) Produced under an FAA production approval;
    (3) A standard part (such as a nut or bolt) manufactured in 
compliance with a government or established industry specification;
    (4) A commercial part as defined in Sec. 21.1 of this part;
    (5) Produced by an owner or operator for maintaining or altering 
that owner or operator's product; or
    (6) Fabricated by an appropriately rated certificate holder with a 
quality system, and consumed in the repair or alteration of a product or 
article in accordance with part 43 of this chapter.
    (b) Except as provided in paragraphs (a)(1) through (a)(2) of this 
section, a person who produces a replacement or modification article for 
sale may not represent that part as suitable for installation on a type-
certificated product.
    (c) Except as provided in paragraphs (a)(1) through (a)(2) of this 
section, a person may not sell or represent an article as suitable for 
installation on an aircraft type-certificated under Sec. Sec. 
21.25(a)(2) or 21.27 unless that article--
    (1) Was declared surplus by the U.S. Armed Forces, and
    (2) Was intended for use on that aircraft model by the U.S. Armed 
Forces.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53385, Oct. 16, 2009; Amdt. 
21-92A, 75 FR 9095, Mar. 1, 2010]



                       Subpart B_Type Certificates

    Source: Docket No. 5085, 29 FR 14564, Oct. 24, 1964, unless 
otherwise noted.



Sec. 21.11  Applicability.

    This subpart prescribes--
    (a) Procedural requirements for the issue of type certificates for 
aircraft, aircraft engines, and propellers; and
    (b) Rules governing the holders of those certificates.



Sec. 21.13  Eligibility.

    Any interested person may apply for a type certificate.

[Amdt. 21-25, 34 FR 14068, Sept. 5, 1969]

[[Page 135]]



Sec. 21.15  Application for type certificate.

    (a) An application for a type certificate is made on a form and in a 
manner prescribed by the FAA and is submitted to the appropriate 
aircraft certification office.
    (b) An application for an aircraft type certificate must be 
accompanied by a three-view drawing of that aircraft and available 
preliminary basic data.
    (c) An application for an aircraft engine type certificate must be 
accompanied by a description of the engine design features, the engine 
operating characteristics, and the proposed engine operating 
limitations.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-40, 
39 FR 35459, Oct. 1, 1974; Amdt. 21-67, 54 FR 39291, Sept. 25, 1989; 
Amdt. 21-92, 74 FR 53385, Oct. 16, 2009]



Sec. 21.16  Special conditions.

    If the FAA finds that the airworthiness regulations of this 
subchapter do not contain adequate or appropriate safety standards for 
an aircraft, aircraft engine, or propeller because of a novel or unusual 
design feature of the aircraft, aircraft engine or propeller, he 
prescribes special conditions and amendments thereto for the product. 
The special conditions are issued in accordance with Part 11 of this 
chapter and contain such safety standards for the aircraft, aircraft 
engine or propeller as the FAA finds necessary to establish a level of 
safety equivalent to that established in the regulations.

[Amdt. 21-19, 32 FR 17851, Dec. 13, 1967; as amended by Amdt. 21-51, 45 
FR 60170, Sept. 11, 1980]



Sec. 21.17  Designation of applicable regulations.

    (a) Except as provided in Sec. Sec. 23.2, 25.2, 27.2, 29.2, and in 
parts 26, 34 and 36 of this subchapter, an applicant for a type 
certificate must show that the aircraft, aircraft engine, or propeller 
concerned meets--
    (1) The applicable requirements of this subchapter that are 
effective on the date of application for that certificate unless--
    (i) Otherwise specified by the FAA; or
    (ii) Compliance with later effective amendments is elected or 
required under this section; and
    (2) Any special conditions prescribed by the FAA.
    (b) For special classes of aircraft, including the engines and 
propellers installed thereon (e.g., gliders, airships, and other 
nonconventional aircraft), for which airworthiness standards have not 
been issued under this subchapter, the applicable requirements will be 
the portions of those other airworthiness requirements contained in 
Parts 23, 25, 27, 29, 31, 33, and 35 found by the FAA to be appropriate 
for the aircraft and applicable to a specific type design, or such 
airworthiness criteria as the FAA may find provide an equivalent level 
of safety to those parts.
    (c) An application for type certification of a transport category 
aircraft is effective for 5 years and an application for any other type 
certificate is effective for 3 years, unless an applicant shows at the 
time of application that his product requires a longer period of time 
for design, development, and testing, and the FAA approves a longer 
period.
    (d) In a case where a type certificate has not been issued, or it is 
clear that a type certificate will not be issued, within the time limit 
established under paragraph (c) of this section, the applicant may--
    (1) File a new application for a type certificate and comply with 
all the provisions of paragraph (a) of this section applicable to an 
original application; or
    (2) File for an extension of the original application and comply 
with the applicable airworthiness requirements of this subchapter that 
were effective on a date, to be selected by the applicant, not earlier 
than the date which precedes the date of issue of the type certificate 
by the time limit established under paragraph (c) of this section for 
the original application.
    (e) If an applicant elects to comply with an amendment to this 
subchapter that is effective after the filing of the application for a 
type certificate, he must also comply with any other amendment that the 
FAA finds is directly related.

[[Page 136]]

    (f) For primary category aircraft, the requirements are:
    (1) The applicable airworthiness requirements contained in parts 23, 
27, 31, 33, and 35 of this subchapter, or such other airworthiness 
criteria as the FAA may find appropriate and applicable to the specific 
design and intended use and provide a level of safety acceptable to the 
FAA.
    (2) The noise standards of part 36 applicable to primary category 
aircraft.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-19, 
32 FR 17851, Dec. 13, 1967; Amdt. 21-24, 34 FR 364, Jan. 10, 1969; Amdt. 
21-42, 40 FR 1033, Jan. 6, 1975; Amdt. 21-58, 50 FR 46877, Nov. 13, 
1985; Amdt. 21-60, 52 FR 8042, Mar. 13, 1987; Amdt. 21-68, 55 FR 32860, 
Aug. 10, 1990; Amdt. 21-69, 56 FR 41051, Aug. 16, 1991; Amdt. 21-70, 57 
FR 41367, Sept. 9, 1992; Amdt. 21-90, 72 FR 63404, Nov. 8, 2007]



Sec. 21.19  Changes requiring a new type certificate.

    Each person who proposes to change a product must apply for a new 
type certificate if the FAA finds that the proposed change in design, 
power, thrust, or weight is so extensive that a substantially complete 
investigation of compliance with the applicable regulations is required.

[Doc. No. 28903, 65 FR 36265, June 7, 2000]



Sec. 21.20  Compliance with applicable requirements.

    The applicant for a type certificate, including an amended or 
supplemental type certificate, must--
    (a) Show compliance with all applicable requirements and must 
provide the FAA the means by which such compliance has been shown; and
    (b) Provide a statement certifying that the applicant has complied 
with the applicable requirements.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53385, Oct. 16, 2009]



Sec. 21.21  Issue of type certificate: normal, utility, acrobatic,
commuter, and transport category aircraft; manned free balloons;
special classes of aircraft; aircraft engines; propellers.

    An applicant is entitled to a type certificate for an aircraft in 
the normal, utility, acrobatic, commuter, or transport category, or for 
a manned free balloon, special class of aircraft, or an aircraft engine 
or propeller, if--
    (a) The product qualifies under Sec. 21.27; or
    (b) The applicant submits the type design, test reports, and 
computations necessary to show that the product to be certificated meets 
the applicable airworthiness, aircraft noise, fuel venting, and exhaust 
emission requirements of this subchapter and any special conditions 
prescribed by the FAA, and the FAA finds--
    (1) Upon examination of the type design, and after completing all 
tests and inspections, that the type design and the product meet the 
applicable noise, fuel venting, and emissions requirements of this 
subchapter, and further finds that they meet the applicable 
airworthiness requirements of this subchapter or that any airworthiness 
provisions not complied with are compensated for by factors that provide 
an equivalent level of safety; and
    (2) For an aircraft, that no feature or characteristic makes it 
unsafe for the category in which certification is requested.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-15, 
32 FR 3735, Mar. 4, 1967; Amdt. 21-27, 34 FR 18368, Nov. 18, 1969; Amdt. 
21-60, 52 FR 8042, Mar. 13, 1987; Amdt. 21-68, 55 FR 32860, Aug. 10, 
1990; Amdt. 21-92, 74 FR 53385, Oct. 16, 2009]



Sec. 21.23  [Reserved]



Sec. 21.24  Issuance of type certificate: primary category aircraft.

    (a) The applicant is entitled to a type certificate for an aircraft 
in the primary category if--
    (1) The aircraft--
    (i) Is unpowered; is an airplane powered by a single, naturally 
aspirated engine with a 61-knot or less Vso stall speed as 
defined in Sec. 23.49; or is a rotorcraft with a 6-pound per square 
foot main rotor disc loading limitation, under sea level standard day 
conditions;
    (ii) Weighs not more than 2,700 pounds; or, for seaplanes, not more 
than 3,375 pounds;
    (iii) Has a maximum seating capacity of not more than four persons, 
including the pilot; and
    (iv) Has an unpressurized cabin.
    (2) The applicant has submitted--

[[Page 137]]

    (i) Except as provided by paragraph (c) of this section, a 
statement, in a form and manner acceptable to the FAA, certifying that: 
the applicant has completed the engineering analysis necessary to 
demonstrate compliance with the applicable airworthiness requirements; 
the applicant has conducted appropriate flight, structural, propulsion, 
and systems tests necessary to show that the aircraft, its components, 
and its equipment are reliable and function properly; the type design 
complies with the airworthiness standards and noise requirements 
established for the aircraft under Sec. 21.17(f); and no feature or 
characteristic makes it unsafe for its intended use;
    (ii) The flight manual required by Sec. 21.5(b), including any 
information required to be furnished by the applicable airworthiness 
standards;
    (iii) Instructions for continued airworthiness in accordance with 
Sec. 21.50(b); and
    (iv) A report that: summarizes how compliance with each provision of 
the type certification basis was determined; lists the specific 
documents in which the type certification data information is provided; 
lists all necessary drawings and documents used to define the type 
design; and lists all the engineering reports on tests and computations 
that the applicant must retain and make available under Sec. 21.49 to 
substantiate compliance with the applicable airworthiness standards.
    (3) The FAA finds that--
    (i) The aircraft complies with those applicable airworthiness 
requirements approved under Sec. 21.17(f) of this part; and
    (ii) The aircraft has no feature or characteristic that makes it 
unsafe for its intended use.
    (b) An applicant may include a special inspection and preventive 
maintenance program as part of the aircraft's type design or 
supplemental type design.
    (c) For aircraft manufactured outside of the United States in a 
country with which the United States has a bilateral airworthiness 
agreement for the acceptance of these aircraft, and from which the 
aircraft is to be imported into the United States--
    (1) The statement required by paragraph (a)(2)(i) of this section 
must be made by the civil airworthiness authority of the exporting 
country; and
    (2) The required manuals, placards, listings, instrument markings, 
and documents required by paragraphs (a) and (b) of this section must be 
submitted in English.

[Doc. No. 23345, 57 FR 41367, Sept. 9, 1992; as amended by Amdt. 21-75, 
62 FR 62808, Nov. 25, 1997]



Sec. 21.25  Issue of type certificate: Restricted category aircraft.

    (a) An applicant is entitled to a type certificate for an aircraft 
in the restricted category for special purpose operations if he shows 
compliance with the applicable noise requirements of Part 36 of this 
chapter, and if he shows that no feature or characteristic of the 
aircraft makes it unsafe when it is operated under the limitations 
prescribed for its intended use, and that the aircraft--
    (1) Meets the airworthiness requirements of an aircraft category 
except those requirements that the FAA finds inappropriate for the 
special purpose for which the aircraft is to be used; or
    (2) Is of a type that has been manufactured in accordance with the 
requirements of and accepted for use by, an Armed Force of the United 
States and has been later modified for a special purpose.
    (b) For the purposes of this section, ``special purpose operations'' 
includes--
    (1) Agricultural (spraying, dusting, and seeding, and livestock and 
predatory animal control);
    (2) Forest and wildlife conservation;
    (3) Aerial surveying (photography, mapping, and oil and mineral 
exploration);
    (4) Patrolling (pipelines, power lines, and canals);
    (5) Weather control (cloud seeding);
    (6) Aerial advertising (skywriting, banner towing, airborne signs 
and public address systems); and
    (7) Any other operation specified by the FAA.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-42, 
40 FR 1033, Jan. 6, 1975]

[[Page 138]]



Sec. 21.27  Issue of type certificate: surplus aircraft of the Armed
Forces.

    (a) Except as provided in paragraph (b) of this section an applicant 
is entitled to a type certificate for an aircraft in the normal, 
utility, acrobatic, commuter, or transport category that was designed 
and constructed in the United States, accepted for operational use, and 
declared surplus by, an Armed Force of the United States, and that is 
shown to comply with the applicable certification requirements in 
paragraph (f) of this section.
    (b) An applicant is entitled to a type certificate for a surplus 
aircraft of the Armed Forces of the United States that is a counterpart 
of a previously type certificated civil aircraft, if he shows compliance 
with the regulations governing the original civil aircraft type 
certificate.
    (c) Aircraft engines, propellers, and their related accessories 
installed in surplus Armed Forces aircraft, for which a type certificate 
is sought under this section, will be approved for use on those aircraft 
if the applicant shows that on the basis of the previous military 
qualifications, acceptance, and service record, the product provides 
substantially the same level of airworthiness as would be provided if 
the engines or propellers were type certificated under Part 33 or 35 of 
this subchapter.
    (d) The FAA may relieve an applicant from strict compliance with a 
specific provision of the applicable requirements in paragraph (f) of 
this section, if the FAA finds that the method of compliance proposed by 
the applicant provides substantially the same level of airworthiness and 
that strict compliance with those regulations would impose a severe 
burden on the applicant. The FAA may use experience that was 
satisfactory to an Armed Force of the United States in making such a 
determination.
    (e) The FAA may require an applicant to comply with special 
conditions and later requirements than those in paragraphs (c) and (f) 
of this section, if the FAA finds that compliance with the listed 
regulations would not ensure an adequate level of airworthiness for the 
aircraft.
    (f) Except as provided in paragraphs (b) through (e) of this 
section, an applicant for a type certificate under this section must 
comply with the appropriate regulations listed in the following table:

------------------------------------------------------------------------
                                 Date accepted for
                                  operational use
        Type of aircraft            by the Armed      Regulations that
                                   Forces of the          apply \1\
                                   United States
------------------------------------------------------------------------
Small reciprocating-engine       Before May 16,     CAR Part 3, as
 powered airplanes.               1956.              effective May 15,
                                 After May 15,       1956.
                                  1956.             CAR Part 3, or 14
                                                     CFR Part 23.
Small turbine engine-powered     Before Oct. 2,     CAR Part 3, as
 airplanes.                       1959.              effective Oct. 1,
                                 After Oct. 1,       1959.
                                  1959.             CAR Part 3 or 14 CFR
                                                     Part 23.
Commuter category airplanes....  After (Feb. 17,
                                  1987).
                                 FAR Part 23 as of
                                  (Feb. 17, 1987)..
Large reciprocating-engine       Before Aug. 26,    CAR Part 4b, as
 powered airplanes.               1955.              effective Aug. 25,
                                 After Aug. 25,      1955.
                                  1955.             CAR Part 4b or 14
                                                     CFR Part 25.
Large turbine engine-powered     Before Oct. 2,     CAR Part 4b, as
 airplanes.                       1959.              effective Oct. 1,
                                 After Oct. 1,       1959.
                                  1959.             CAR Part 4b or 14
                                                     CFR Part 25.
Rotorcraft with maximum
 certificated takeoff weight
 of:
  6,000 pounds or less.........  Before Oct. 2,     CAR Part 6, as
                                  1959.              effective Oct. 1,
                                 After Oct. 1,       1959.
                                  1959.             CAR Part 6, or 14
                                                     CFR Part 27.
  Over 6,000 pounds............  Before Oct. 2,     CAR Part 7, as
                                  1959.              effective Oct. 1,
                                 After Oct. 1,       1959.
                                  1959.             CAR Part 7, or 14
                                                     CFR Part 29.
------------------------------------------------------------------------
\1\ Where no specific date is listed, the applicable regulations are
  those in effect on the date that the first aircraft of the particular
  model was accepted for operational use by the Armed Forces.


[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-59, 
52 FR 1835, Jan. 15, 1987; 52 FR 7262, Mar. 9, 1987; 70 FR 2325, Jan. 
13, 2005; Amdt. 21-92, 74 FR 53386, Oct. 16, 2009]



Sec. 21.29  Issue of type certificate: import products.

    (a) The FAA may issue a type certificate for a product that is 
manufactured in a foreign country or jurisdiction with which the United 
States has an agreement for the acceptance of these products for export 
and import and

[[Page 139]]

that is to be imported into the United States if--
    (1) The applicable State of Design certifies that the product has 
been examined, tested, and found to meet--
    (i) The applicable aircraft noise, fuel venting, and exhaust 
emissions requirements of this subchapter as designated in Sec. 21.17, 
or the applicable aircraft noise, fuel venting, and exhaust emissions 
requirements of the State of Design, and any other requirements the FAA 
may prescribe to provide noise, fuel venting, and exhaust emission 
levels no greater than those provided by the applicable aircraft noise, 
fuel venting, and exhaust emission requirements of this subchapter as 
designated in Sec. 21.17; and
    (ii) The applicable airworthiness requirements of this subchapter as 
designated in Sec. 21.17, or the applicable airworthiness requirements 
of the State of Design and any other requirements the FAA may prescribe 
to provide a level of safety equivalent to that provided by the 
applicable airworthiness requirements of this subchapter as designated 
in Sec. 21.17;
    (2) The applicant has provided technical data to show the product 
meets the requirements of paragraph (a)(1) of this section; and
    (3) The manuals, placards, listings, and instrument markings 
required by the applicable airworthiness (and noise, where applicable) 
requirements are presented in the English language.
    (b) A product type certificated under this section is considered to 
be type certificated under the noise standards of part 36 of this 
subchapter and the fuel venting and exhaust emission standards of part 
34 of this subchapter. Compliance with parts 36 and 34 of this 
subchapter is certified under paragraph (a)(1)(i) of this section, and 
the applicable airworthiness standards of this subchapter, or an 
equivalent level of safety, with which compliance is certified under 
paragraph (a)(1)(ii) of this section.

[Amdt. 21-92, 74 FR 53386, Oct. 16, 2009]



Sec. 21.31  Type design.

    The type design consists of--
    (a) The drawings and specifications, and a listing of those drawings 
and specifications, necessary to define the configuration and the design 
features of the product shown to comply with the requirements of that 
part of this subchapter applicable to the product;
    (b) Information on dimensions, materials, and processes necessary to 
define the structural strength of the product;
    (c) The Airworthiness Limitations section of the Instructions for 
Continued Airworthiness as required by parts 23, 25, 26, 27, 29, 31, 33 
and 35 of this subchapter, or as otherwise required by the FAA; and as 
specified in the applicable airworthiness criteria for special classes 
of aircraft defined in Sec. 21.17(b); and
    (d) For primary category aircraft, if desired, a special inspection 
and preventive maintenance program designed to be accomplished by an 
appropriately rated and trained pilot-owner.
    (e) Any other data necessary to allow, by comparison, the 
determination of the airworthiness, noise characteristics, fuel venting, 
and exhaust emissions (where applicable) of later products of the same 
type.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-27, 
34 FR 18363, Nov. 18, 1969; Amdt. 21-51, 45 FR 60170, Sept. 11, 1980; 
Amdt. 21-60, 52 FR 8042, Mar. 13, 1987; Amdt. 21-68, 55 FR 32860, Aug. 
10, 1990; Amdt. 21-70, 57 FR 41368, Sept. 9, 1992; Amdt. 21-90, 72 FR 
63404, Nov. 8, 2007]



Sec. 21.33  Inspection and tests.

    (a) Each applicant must allow the FAA to make any inspection and any 
flight and ground test necessary to determine compliance with the 
applicable requirements of this subchapter. However, unless otherwise 
authorized by the FAA--
    (1) No aircraft, aircraft engine, propeller, or part thereof may be 
presented to the FAA for test unless compliance with paragraphs (b)(2) 
through (b)(4) of this section has been shown for that aircraft, 
aircraft engine, propeller, or part thereof; and
    (2) No change may be made to an aircraft, aircraft engine, 
propeller, or part thereof between the time that compliance with 
paragraphs (b)(2) through (b)(4) of this section is shown for that 
aircraft, aircraft engine, propeller, or part thereof and the time that 
it is presented to the FAA for test.

[[Page 140]]

    (b) Each applicant must make all inspections and tests necessary to 
determine--
    (1) Compliance with the applicable airworthiness, aircraft noise, 
fuel venting, and exhaust emission requirements;
    (2) That materials and products conform to the specifications in the 
type design;
    (3) That parts of the products conform to the drawings in the type 
design; and
    (4) That the manufacturing processes, construction and assembly 
conform to those specified in the type design.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-17, 
32 FR 14926, Oct. 28, 1967; Amdt. 21-27, 34 FR 18363, Nov. 18, 1969; 
Amdt. 21-44, 41 FR 55463, Dec. 20, 1976; Amdt. 21-68, 55 FR 32860, Aug. 
10, 1990; Amdt. 21-68, 55 FR 32860, Aug. 10, 1990; Amdt. 21-92, 74 FR 
53386, Oct. 16, 2009]



Sec. 21.35  Flight tests.

    (a) Each applicant for an aircraft type certificate (other than 
under Sec. Sec. 21.24 through 21.29) must make the tests listed in 
paragraph (b) of this section. Before making the tests the applicant 
must show--
    (1) Compliance with the applicable structural requirements of this 
subchapter;
    (2) Completion of necessary ground inspections and tests;
    (3) That the aircraft conforms with the type design; and
    (4) That the FAA received a flight test report from the applicant 
(signed, in the case of aircraft to be certificated under Part 25 [New] 
of this chapter, by the applicant's test pilot) containing the results 
of his tests.
    (b) Upon showing compliance with paragraph (a) of this section, the 
applicant must make all flight tests that the FAA finds necessary--
    (1) To determine compliance with the applicable requirements of this 
subchapter; and
    (2) For aircraft to be certificated under this subchapter, except 
gliders and except reciprocating engine powered airplanes of 6,000 lbs. 
or less maximum certificated weight that are to be certificated under 
part 23 of this chapter, to determine whether there is reasonable 
assurance that the aircraft, its components, and its equipment are 
reliable and function properly.
    (c) Each applicant must, if practicable, make the tests prescribed 
in paragraph (b)(2) of this section upon the aircraft that was used to 
show compliance with--
    (1) Paragraph (b)(1) of this section; and
    (2) For rotorcraft, the rotor drive endurance tests prescribed in 
Sec. 27.923 or Sec. 29.923 of this chapter, as applicable.
    (d) Each applicant must show for each flight test (except in a 
glider or a manned free balloon) that adequate provision is made for the 
flight test crew for emergency egress and the use of parachutes.
    (e) Except in gliders and manned free balloons, an applicant must 
discontinue flight tests under this section until he shows that 
corrective action has been taken, whenever--
    (1) The applicant's test pilot is unable or unwilling to make any of 
the required flight tests; or
    (2) Items of noncompliance with requirements are found that may make 
additional test data meaningless or that would make further testing 
unduly hazardous.
    (f) The flight tests prescribed in paragraph (b)(2) of this section 
must include--
    (1) For aircraft incorporating turbine engines of a type not 
previously used in a type certificated aircraft, at least 300 hours of 
operation with a full complement of engines that conform to a type 
certificate; and
    (2) For all other aircraft, at least 150 hours of operation.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-40, 
39 FR 35459, Oct. 1, 1974; Amdt. 21-51, 45 FR 60170, Sept. 11, 1980; 
Amdt. 21-70, 57 FR 41368, Sept. 9, 1992; Amdt. 21-95, 76 FR 64233, Oct. 
18, 2011]



Sec. 21.37  Flight test pilot.

    Each applicant for a normal, utility, acrobatic, commuter, or 
transport category aircraft type certificate must provide a person 
holding an appropriate pilot certificate to make the flight tests 
required by this part.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-59, 
52 FR 1835, Jan. 15, 1987]

[[Page 141]]



Sec. 21.39  Flight test instrument calibration and correction report.

    (a) Each applicant for a normal, utility, acrobatic, commuter, or 
transport category aircraft type certificate must submit a report to the 
FAA showing the computations and tests required in connection with the 
calibration of instruments used for test purposes and in the correction 
of test results to standard atmospheric conditions.
    (b) Each applicant must allow the FAA to conduct any flight tests 
that he finds necessary to check the accuracy of the report submitted 
under paragraph (a) of this section.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-59, 
52 FR 1835, Jan. 15, 1987]



Sec. 21.41  Type certificate.

    Each type certificate is considered to include the type design, the 
operating limitations, the certificate data sheet, the applicable 
regulations of this subchapter with which the FAA records compliance, 
and any other conditions or limitations prescribed for the product in 
this subchapter.



Sec. 21.43  Location of manufacturing facilities.

    Except as provided in Sec. 21.29, the FAA does not issue a type 
certificate if the manufacturing facilities for the product are located 
outside of the United States, unless the FAA finds that the location of 
the manufacturer's facilities places no undue burden on the FAA in 
administering applicable airworthiness requirements.



Sec. 21.45  Privileges.

    The holder or licensee of a type certificate for a product may--
    (a) In the case of aircraft, upon compliance with Sec. Sec. 21.173 
through 21.189, obtain airworthiness certificates;
    (b) In the case of aircraft engines or propellers, obtain approval 
for installation on certificated aircraft;
    (c) In the case of any product, upon compliance with subpart G of 
this part, obtain a production certificate for the type certificated 
product;
    (d) Obtain approval of replacement parts for that product.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Amdt. 21-92, 
74 FR 53386, Oct. 16, 2009]



Sec. 21.47  Transferability.

    (a) A holder of a type certificate may transfer it or make it 
available to other persons by licensing agreements.
    (b) For a type certificate transfer in which the State of Design 
will remain the same, each transferor must, before such a transfer, 
notify in writing the appropriate aircraft certification office. This 
notification must include the applicable type certificate number, the 
name and address of the transferee, and the anticipated date of the 
transfer.
    (c) For a type certificate transfer in which the State of Design is 
changing, a type certificate may only be transferred to or from a person 
subject to the authority of another State of Design if the United States 
has an agreement with that State of Design for the acceptance of the 
affected product for export and import. Each transferor must notify the 
appropriate aircraft certification office before such a transfer in a 
form and manner acceptable to the FAA. This notification must include 
the applicable type certificate number; the name, address, and country 
of residence of the transferee; and the anticipated date of the 
transfer.
    (d) Before executing or terminating a licensing agreement that makes 
a type certificate available to another person, the type certificate 
holder must notify in writing the appropriate aircraft certification 
office. This notification must include the type certificate number 
addressed by the licensing agreement, the name and address of the 
licensee, the extent of authority granted the licensee, and the 
anticipated date of the agreement.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53386, Oct. 16, 2009]



Sec. 21.49  Availability.

    The holder of a type certificate must make the certificate available 
for examination upon the request of the FAA

[[Page 142]]

or the National Transportation Safety Board.

[Doc. No. 5085, 29 FR 14564, Oct. 24, 1964, as amended by Docket No. 
8084, 32 FR 5769, Apr. 11, 1967]



Sec. 21.50  Instructions for continued airworthiness and manufacturer's
maintenance manuals having airworthiness limitations sections.

    (a) The holder of a type certificate for a rotorcraft for which a 
Rotorcraft Maintenance Manual containing an ``Airworthiness 
Limitations'' section has been issued under Sec. 27.1529 (a)(2) or 
Sec. 29.1529 (a)(2) of this chapter, and who obtains approval of 
changes to any replacement time, inspection interval, or related 
procedure in that section of the manual, must make those changes 
available upon request to any operator of the same type of rotorcraft.
    (b) The holder of a design approval, including either the type 
certificate or supplemental type certificate for an aircraft, aircraft 
engine, or propeller for which application was made after January 28, 
1981, must furnish at least one set of complete Instructions for 
Continued Airworthiness to the owner of each type aircraft, aircraft 
engine, or propeller upon its delivery, or upon issuance of the first 
standard airworthiness certificate for the affected aircraft, whichever 
occurs later. The Instructions must be prepared in accordance with 
Sec. Sec. 23.1529, 25.1529, 25.1729, 27.1529, 29.1529, 31.82, 33.4, 
35.4, or part 26 of this subchapter, or as specified in the applicable 
airworthiness criteria for special classes of aircraft defined in Sec. 
21.17(b), as applicable. If the holder of a design approval chooses to 
designate parts as commercial, it must include in the Instructions for 
Continued Airworthiness a list of commercial parts submitted in 
accordance with the provisions of paragraph (c) of this section. 
Thereafter, the holder of a design approval must make those instructions 
available to any other person required by this chapter to comply with 
any of the terms of those instructions. In addition, changes to the 
Instructions for Continued Airworthiness shall be made available to any 
person required by this chapter to comply with any of those 
instructions.
    (c) To designate commercial parts, the holder of a design approval, 
in a manner acceptable to the FAA, must submit:
    (1) A Commercial Parts List;
    (2) Data for each part on the List showing that:
    (i) The failure of the commercial part, as installed in the product, 
would not degrade the level of safety of the product; and
    (ii) The part is produced only under the commercial part 
manufacturer's specification and marked only with the commercial part 
manufacturer's markings; and
    (3) Any other data necessary for the FAA to approve the List.

[Amdt. 21-23, 33 FR 14105, Sept. 18, 1968, as amended by Amdt. 21-51, 45 
FR 60170, Sept. 11, 1980; Amdt. 21-60, 52 FR 8042, Mar. 13, 1987; Amdt. 
21-90, 72 FR 63404, Nov. 8, 2007; Amdt. 21-92, 74 FR 53386, Oct. 16, 
2009]



Sec. 21.51  Duration.

    A type certificate is effective until surrendered, suspended, 
revoked, or a termination date is otherwise established by the FAA.



Sec. 21.53  Statement of conformity.

    (a) Each applicant must provide, in a form and manner acceptable to 
the FAA, a statement that each aircraft engine or propeller presented 
for type certification conforms to its type design.
    (b) Each applicant must submit a statement of conformity to the FAA 
for each aircraft or part thereof presented to the FAA for tests. This 
statement of conformity must include a statement that the applicant has 
complied with Sec. 21.33(a) (unless otherwise authorized under that 
paragraph).

[Amdt. 21-17, 32 FR 14926, Oct. 28, 1967, as amended by Amdt. 21-92, 74 
FR 53386, Oct. 16, 2009]



Sec. 21.55  Responsibility of type certificate holders to provide
written licensing agreements.

    A type certificate holder who allows a person to use the type 
certificate to manufacture a new aircraft, aircraft engine, or propeller 
must provide that

[[Page 143]]

person with a written licensing agreement acceptable to the FAA.

[Doc. No. FAA-2003-14825, 71 FR 52258, Sept. 1, 2006]



                 Subpart C_Provisional Type Certificates

    Source: Docket No. 5085, 29 FR 14566, Oct. 24, 1964, unless 
otherwise noted.



Sec. 21.71  Applicability.

    This subpart prescribes--
    (a) Procedural requirements for the issue of provisional type 
certificates, amendments to provisional type certificates, and 
provisional amendments to type certificates; and
    (b) Rules governing the holders of those certificates.



Sec. 21.73  Eligibility.

    (a) Any manufacturer of aircraft manufactured within the United 
States who is a United States citizen may apply for Class I or Class II 
provisional type certificates, for amendments to provisional type 
certificates held by him, and for provisional amendments to type 
certificates held by him.
    (b) Any manufacturer of aircraft in a State of Manufacture subject 
to the provisions of an agreement with the United States for the 
acceptance of those aircraft for export and import may apply for a Class 
II provisional type certificate, for amendments to provisional type 
certificates held by him, and for provisional amendments to type 
certificates held by him.
    (c) An aircraft engine manufacturer who is a United States citizen 
and who has altered a type certificated aircraft by installing different 
type certificated aircraft engines manufactured by him within the United 
States may apply for a Class I provisional type certificate for the 
aircraft, and for amendments to Class I provisional type certificates 
held by him, if the basic aircraft, before alteration, was type 
certificated in the normal, utility, acrobatic, commuter, or transport 
category.

[Doc. No. 5085, 29 FR 14566, Oct. 24, 1964, as amended by Amdt. 21-12, 
31 FR 13380, Oct. 15, 1966; Amdt. 21-59, 52 FR 1836, Jan. 15, 1987; 
Amdt. 21-92, 74 FR 53387, Oct. 16, 2009]



Sec. 21.75  Application.

    Each applicant for a provisional type certificate, for an amendment 
thereto, or for a provisional amendment to a type certificate must apply 
to the appropriate aircraft certification office and provide the 
information required by this subpart.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53387, Oct. 16, 2009]



Sec. 21.77  Duration.

    (a) Unless sooner surrendered, superseded, revoked, or otherwise 
terminated, provisional type certificates and amendments thereto are 
effective for the periods specified in this section.
    (b) A Class I provisional type certificate is effective for 24 
months after the date of issue.
    (c) A Class II provisional type certificate is effective for twelve 
months after the date of issue.
    (d) An amendment to a Class I or Class II provisional type 
certificate is effective for the duration of the amended certificate.
    (e) A provisional amendment to a type certificate is effective for 
six months after its approval or until the amendment of the type 
certificate is approved, whichever is first.

[Doc. No. 5085, 29 FR 14566, Oct. 24, 1964 as amended by Amdt. 21-7, 30 
FR 14311, Nov. 16, 1965]



Sec. 21.79  Transferability.

    Provisional type certificates are not transferable.



Sec. 21.81  Requirements for issue and amendment of Class I provisional
type certificates.

    (a) An applicant is entitled to the issue or amendment of a Class I 
provisional type certificate if he shows compliance with this section 
and the FAA finds that there is no feature, characteristic, or condition 
that would make the aircraft unsafe when operated in accordance with the 
limitations established in paragraph (e) of this section and in Sec. 
91.317 of this chapter.
    (b) The applicant must apply for the issue of a type or supplemental 
type certificate for the aircraft.
    (c) The applicant must certify that--

[[Page 144]]

    (1) The aircraft has been designed and constructed in accordance 
with the airworthiness requirements applicable to the issue of the type 
or supplemental type certificate applied for;
    (2) The aircraft substantially meets the applicable flight 
characteristic requirements for the type or supplemental type 
certificate applied for; and
    (3) The aircraft can be operated safely under the appropriate 
operating limitations specified in paragraph (a) of this section.
    (d) The applicant must submit a report showing that the aircraft had 
been flown in all maneuvers necessary to show compliance with the flight 
requirements for the issue of the type or supplemental type certificate 
applied for, and to establish that the aircraft can be operated safely 
in accordance with the limitations contained in this subchapter.
    (e) The applicant must establish all limitations required for the 
issue of the type or supplemental type certificate applied for, 
including limitations on weights, speeds, flight maneuvers, loading, and 
operation of controls and equipment unless, for each limitation not so 
established, appropriate operating restrictions are established for the 
aircraft.
    (f) The applicant must establish an inspection and maintenance 
program for the continued airworthiness of the aircraft.
    (g) The applicant must show that a prototype aircraft has been flown 
for at least 50 hours under an experimental certificate issued under 
Sec. Sec. 21.191 through 21.195, or under the auspices of an Armed 
Force of the United States. However, in the case of an amendment to a 
provisional type certificate, the FAA may reduce the number of required 
flight hours.

[Doc. No. 5085, 29 FR 14566, Oct. 24, 1964, as amended by Amdt. 21-66, 
54 FR 34329, Aug. 18, 1989]



Sec. 21.83  Requirements for issue and amendment of Class II provisional
type certificates.

    (a) An applicant who manufactures aircraft within the United States 
is entitled to the issue or amendment of a Class II provisional type 
certificate if he shows compliance with this section and the FAA finds 
that there is no feature, characteristic, or condition that would make 
the aircraft unsafe when operated in accordance with the limitations in 
paragraph (h) of this section, and Sec. Sec. 91.317 and 121.207 of this 
chapter.
    (b) An applicant who manufactures aircraft in a country with which 
the United States has an agreement for the acceptance of those aircraft 
for export and import is entitled to the issue or amendment of a Class 
II provisional type certificate if the country in which the aircraft was 
manufactured certifies that the applicant has shown compliance with this 
section, that the aircraft meets the requirements of paragraph (f) of 
this section and that there is no feature, characteristic, or condition 
that would make the aircraft unsafe when operated in accordance with the 
limitations in paragraph (h) of this section and Sec. Sec. 91.317 and 
121.207 of this chapter.
    (c) The applicant must apply for a type certificate, in the 
transport category, for the aircraft.
    (d) The applicant must hold a U.S. type certificate for at least one 
other aircraft in the same transport category as the subject aircraft.
    (e) The FAA's official flight test program or the flight test 
program conducted by the authorities of the country in which the 
aircraft was manufactured, with respect to the issue of a type 
certificate for that aircraft, must be in progress.
    (f) The applicant or, in the case of a foreign manufactured 
aircraft, the country in which the aircraft was manufactured, must 
certify that--
    (1) The aircraft has been designed and constructed in accordance 
with the airworthiness requirements applicable to the issue of the type 
certificate applied for;
    (2) The aircraft substantially complies with the applicable flight 
characteristic requirements for the type certificate applied for; and
    (3) The aircraft can be operated safely under the appropriate 
operating limitations in this subchapter.
    (g) The applicant must submit a report showing that the aircraft has 
been flown in all maneuvers necessary to show compliance with the flight 
requirements for the issue of the type

[[Page 145]]

certificate and to establish that the aircraft can be operated safely in 
accordance with the limitations in this subchapter.
    (h) The applicant must prepare a provisional aircraft flight manual 
containing all limitations required for the issue of the type 
certificate applied for, including limitations on weights, speeds, 
flight maneuvers, loading, and operation of controls and equipment 
unless, for each limitation not so established, appropriate operating 
restrictions are established for the aircraft.
    (i) The applicant must establish an inspection and maintenance 
program for the continued airworthiness of the aircraft.
    (j) The applicant must show that a prototype aircraft has been flown 
for at least 100 hours. In the case of an amendment to a provisional 
type certificate, the FAA may reduce the number of required flight 
hours.

[Amdt. 21-12, 31 FR 13386, Oct. 15, 1966, as amended by Amdt. 21-66, 54 
FR 34329, Aug. 18, 1989]



Sec. 21.85  Provisional amendments to type certificates.

    (a) An applicant who manufactures aircraft within the United States 
is entitled to a provisional amendment to a type certificate if he shows 
compliance with this section and the FAA finds that there is no feature, 
characteristic, or condition that would make the aircraft unsafe when 
operated under the appropriate limitations contained in this subchapter.
    (b) An applicant who manufactures aircraft in a foreign country with 
which the United States has an agreement for the acceptance of those 
aircraft for export and import is entitled to a provisional amendment to 
a type certificate if the country in which the aircraft was manufactured 
certifies that the applicant has shown compliance with this section, 
that the aircraft meets the requirements of paragraph (e) of this 
section and that there is no feature, characteristic, or condition that 
would make the aircraft unsafe when operated under the appropriate 
limitations contained in this subchapter.
    (c) The applicant must apply for an amendment to the type 
certificate.
    (d) The FAA's official flight test program or the flight test 
program conducted by the authorities of the country in which the 
aircraft was manufactured, with respect to the amendment of the type 
certificate, must be in progress.
    (e) The applicant or, in the case of foreign manufactured aircraft, 
the country in which the aircraft was manufactured, must certify that--
    (1) The modification involved in the amendment to the type 
certificate has been designed and constructed in accordance with the 
airworthiness requirements applicable to the issue of the type 
certificate for the aircraft;
    (2) The aircraft substantially complies with the applicable flight 
characteristic requirements for the type certificate; and
    (3) The aircraft can be operated safely under the appropriate 
operating limitations in this subchapter.
    (f) The applicant must submit a report showing that the aircraft 
incorporating the modifications involved has been flown in all maneuvers 
necessary to show compliance with the flight requirements applicable to 
those modifications and to establish that the aircraft can be operated 
safely in accordance with the limitations specified in Sec. Sec. 91.317 
and 121.207 of this chapter.
    (g) The applicant must establish and publish, in a provisional 
aircraft flight manual or other document and on appropriate placards, 
all limitations required for the issue of the type certificate applied 
for, including weight, speed, flight maneuvers, loading, and operation 
of controls and equipment, unless, for each limitation not so 
established, appropriate operating restrictions are established for the 
aircraft.
    (h) The applicant must establish an inspection and maintenance 
program for the continued airworthiness of the aircraft.
    (i) The applicant must operate a prototype aircraft modified in 
accordance with the corresponding amendment to

[[Page 146]]

the type certificate for the number of hours found necessary by the FAA.

[Amdt. 21-12, 31 FR 13388, Oct. 15, 1966, as amended by Amdt. 21-66, 54 
FR 34329, Aug. 18, 1989]



                 Subpart D_Changes to Type Certificates

    Source: Docket No. 5085, 29 FR 14567, Oct. 24, 1964, unless 
otherwise noted.



Sec. 21.91  Applicability.

    This subpart prescribes procedural requirements for the approval of 
changes to type certificates.



Sec. 21.93  Classification of changes in type design.

    (a) In addition to changes in type design specified in paragraph (b) 
of this section, changes in type design are classified as minor and 
major. A ``minor change'' is one that has no appreciable effect on the 
weight, balance, structural strength, reliability, operational 
characteristics, or other characteristics affecting the airworthiness of 
the product. All other changes are ``major changes'' (except as provided 
in paragraph (b) of this section).
    (b) For the purpose of complying with Part 36 of this chapter, and 
except as provided in paragraphs (b)(2), (b)(3), and (b)(4) of this 
section, any voluntary change in the type design of an aircraft that may 
increase the noise levels of that aircraft is an ``acoustical change'' 
(in addition to being a minor or major change as classified in paragraph 
(a) of this section) for the following aircraft:
    (1) Transport category large airplanes.
    (2) Jet (Turbojet powered) airplanes (regardless of category). For 
airplanes to which this paragraph applies, ``acoustical changes'' do not 
include changes in type design that are limited to one of the 
following--
    (i) Gear down flight with one or more retractable landing gear down 
during the entire flight, or
    (ii) Spare engine and nacelle carriage external to the skin of the 
airplane (and return of the pylon or other external mount), or
    (iii) Time-limited engine and/or nacelle changes, where the change 
in type design specifies that the airplane may not be operated for a 
period of more than 90 days unless compliance with the applicable 
acoustical change provisions of Part 36 of this chapter is shown for 
that change in type design.
    (3) Propeller driven commuter category and small airplanes in the 
primary, normal, utility, acrobatic, transport, and restricted 
categories, except for airplanes that are:
    (i) Designated for ``agricultural aircraft operations'' (as defined 
in Sec. 137.3 of this chapter, effective January 1, 1966) to which 
Sec. 36.1583 of this chapter does not apply, or
    (ii) Designated for dispensing fire fighting materials to which 
Sec. 36.1583 of this chapter does not apply, or
    (iii) U.S. registered, and that had flight time prior to January 1, 
1955 or
    (iv) Land configured aircraft reconfigured with floats or skis. This 
reconfiguration does not permit further exception from the requirements 
of this section upon any acoustical change not enumerated in Sec. 
21.93(b).
    (4) Helicopters except:
    (i) Those helicopters that are designated exclusively:
    (A) For ``agricultural aircraft operations'', as defined in Sec. 
137.3 of this chapter, as effective on January 1, 1966;
    (B) For dispensing fire fighting materials; or
    (C) For carrying external loads, as defined in Sec. 133.1(b) of 
this chapter, as effective on December 20, 1976.
    (ii) Those helicopters modified by installation or removal of 
external equipment. For purposes of this paragraph, ``external 
equipment'' means any instrument, mechanism, part, apparatus, 
appurtenance, or accessory that is attached to, or extends from, the 
helicopter exterior but is not used nor is intended to be used in 
operating or controlling a helicopter in flight and is not part of an 
airframe or engine. An ``acoustical change'' does not include:
    (A) Addition or removal of external equipment;
    (B) Changes in the airframe made to accommodate the addition or 
removal of external equipment, to provide for an external load attaching 
means, to facilitate the use of external equipment or external loads, or 
to facilitate the safe operation of the helicopter

[[Page 147]]

with external equipment mounted to, or external loads carried by, the 
helicopter;
    (C) Reconfiguration of the helicopter by the addition or removal of 
floats and skis;
    (D) Flight with one or more doors and/or windows removed or in an 
open position; or
    (E) Any changes in the operational limitations placed on the 
helicopter as a consequence of the addition or removal of external 
equipment, floats, and skis, or flight operations with doors and/or 
windows removed or in an open position.
    (5) Tiltrotors.
    (c) For purposes of complying with part 34 of this chapter, any 
voluntary change in the type design of the airplane or engine which may 
increase fuel venting or exhaust emissions is an ``emissions change.''

[Amdt. 21-27, 34 FR 18363, Nov. 18, 1969]

    Editorial Note: For Federal Register citations affecting Sec. 
21.93, see the List of CFR Sections Affected, which appears in the 
Finding Aids section of the printed volume and at www.fdsys.gov.



Sec. 21.95  Approval of minor changes in type design.

    Minor changes in a type design may be approved under a method 
acceptable to the FAA before submitting to the FAA any substantiating or 
descriptive data.



Sec. 21.97  Approval of major changes in type design.

    (a) An applicant for approval of a major change in type design 
must--
    (1) Provide substantiating data and necessary descriptive data for 
inclusion in the type design;
    (2) Show that the change and areas affected by the change comply 
with the applicable requirements of this subchapter, and provide the FAA 
the means by which such compliance has been shown; and
    (3) Provide a statement certifying that the applicant has complied 
with the applicable requirements.
    (b) Approval of a major change in the type design of an aircraft 
engine is limited to the specific engine configuration upon which the 
change is made unless the applicant identifies in the necessary 
descriptive data for inclusion in the type design the other 
configurations of the same engine type for which approval is requested 
and shows that the change is compatible with the other configurations.

[Amdt. 21-40, 39 FR 35459, Oct. 1, 1974, as amended by Amdt. 21-92, 74 
FR 53387, Oct. 16, 2009; Amdt. 21-96, 77 FR 71695, Dec. 4, 2012]



Sec. 21.99  Required design changes.

    (a) When an Airworthiness Directive is issued under Part 39 the 
holder of the type certificate for the product concerned must--
    (1) If the FAA finds that design changes are necessary to correct 
the unsafe condition of the product, and upon his request, submit 
appropriate design changes for approval; and
    (2) Upon approval of the design changes, make available the 
descriptive data covering the changes to all operators of products 
previously certificated under the type certificate.
    (b) In a case where there are no current unsafe conditions, but the 
FAA or the holder of the type certificate finds through service 
experience that changes in type design will contribute to the safety of 
the product, the holder of the type certificate may submit appropriate 
design changes for approval. Upon approval of the changes, the 
manufacturer must make information on the design changes available to 
all operators of the same type of product.

[Doc. No. 5085, 29 FR 14567, Oct. 24, 1964, as amended by Amdt. 21-3, 30 
FR 8826, July 24, 1965]



Sec. 21.101  Designation of applicable regulations.

    (a) An applicant for a change to a type certificate must show that 
the change and areas affected by the change comply with the 
airworthiness requirements applicable to the category of the product in 
effect on the date of the application for the change and with parts 34 
and 36 of this chapter. Exceptions are detailed in paragraphs (b) and 
(c) of this section.
    (b) Except as provided in paragraph (g) of this section, if 
paragraphs (b)(1), (2), or (3) of this section apply, an applicant may 
show that the change and areas affected by the change comply

[[Page 148]]

with an earlier amendment of a regulation required by paragraph (a) of 
this section, and of any other regulation the FAA finds is directly 
related. However, the earlier amended regulation may not precede either 
the corresponding regulation incorporated by reference in the type 
certificate, or any regulation in Sec. Sec. 23.2, 25.2, 27.2, or 29.2 
of this subchapter that is related to the change. The applicant may show 
compliance with an earlier amendment of a regulation for any of the 
following:
    (1) A change that the FAA finds not to be significant. In 
determining whether a specific change is significant, the FAA considers 
the change in context with all previous relevant design changes and all 
related revisions to the applicable regulations incorporated in the type 
certificate for the product. Changes that meet one of the following 
criteria are automatically considered significant:
    (i) The general configuration or the principles of construction are 
not retained.
    (ii) The assumptions used for certification of the product to be 
changed do not remain valid.
    (2) Each area, system, component, equipment, or appliance that the 
FAA finds is not affected by the change.
    (3) Each area, system, component, equipment, or appliance that is 
affected by the change, for which the FAA finds that compliance with a 
regulation described in paragraph (a) of this section would not 
contribute materially to the level of safety of the product or would be 
impractical.
    (c) An applicant for a change to an aircraft (other than a 
rotorcraft) of 6,000 pounds or less maximum weight, or to a non-turbine 
rotorcraft of 3,000 pounds or less maximum weight may show that the 
change and areas affected by the change comply with the regulations 
incorporated by reference in the type certificate. However, if the FAA 
finds that the change is significant in an area, the FAA may designate 
compliance with an amendment to the regulation incorporated by reference 
in the type certificate that applies to the change and any regulation 
that the FAA finds is directly related, unless the FAA also finds that 
compliance with that amendment or regulation would not contribute 
materially to the level of safety of the product or would be 
impractical.
    (d) If the FAA finds that the regulations in effect on the date of 
the application for the change do not provide adequate standards with 
respect to the proposed change because of a novel or unusual design 
feature, the applicant must also comply with special conditions, and 
amendments to those special conditions, prescribed under the provisions 
of Sec. 21.16, to provide a level of safety equal to that established 
by the regulations in effect on the date of the application for the 
change.
    (e) An application for a change to a type certificate for a 
transport category aircraft is effective for 5 years, and an application 
for a change to any other type certificate is effective for 3 years. If 
the change has not been approved, or if it is clear that it will not be 
approved under the time limit established under this paragraph, the 
applicant may do either of the following:
    (1) File a new application for a change to the type certificate and 
comply with all the provisions of paragraph (a) of this section 
applicable to an original application for a change.
    (2) File for an extension of the original application and comply 
with the provisions of paragraph (a) of this section. The applicant must 
then select a new application date. The new application date may not 
precede the date the change is approved by more than the time period 
established under this paragraph (e).
    (f) For aircraft certificated under Sec. Sec. 21.17(b), 21.24, 
21.25, and 21.27 the airworthiness requirements applicable to the 
category of the product in effect on the date of the application for the 
change include each airworthiness requirement that the FAA finds to be 
appropriate for the type certification of the aircraft in accordance 
with those sections.
    (g) Notwithstanding paragraph (b) of this section, for transport 
category airplanes, the applicant must show compliance with each 
applicable provision of part 26 of this chapter, unless the applicant 
has elected or was required to comply with a corresponding amendment to 
part 25 of this chapter that

[[Page 149]]

was issued on or after the date of the applicable part 26 provision.

[Doc. No. 28903, 65 FR 36266, June 7, 2000, as amended by Amdt. 21-90, 
72 FR 63404, Nov. 8, 2007; Amdt. 21-96, 77 FR 71695, Dec. 4, 2012]



                Subpart E_Supplemental Type Certificates

    Source: Docket No. 5085, 29 FR 14568, Oct. 24, 1964, unless 
otherwise noted.



Sec. 21.111  Applicability.

    This subpart prescribes procedural requirements for the issue of 
supplemental type certificates.



Sec. 21.113  Requirement for supplemental type certificate.

    (a) If a person holds the TC for a product and alters that product 
by introducing a major change in type design that does not require an 
application for a new TC under Sec. 21.19, that person must either 
apply to the appropriate aircraft certification office for an STC or 
apply to amend the original type certificate under subpart D of this 
part.
    (b) If a person does not hold the TC for a product and alters that 
product by introducing a major change in type design that does not 
require an application for a new TC under Sec. 21.19, that person must 
apply to the appropriate aircraft certification office for an STC.
    (c) The application for an STC must be made in the form and manner 
prescribed by the FAA.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53387, Oct. 16, 2009]



Sec. 21.115  Applicable requirements.

    (a) Each applicant for a supplemental type certificate must show 
that the altered product meets applicable requirements specified in 
Sec. 21.101 and, in the case of an acoustical change described in Sec. 
21.93(b), show compliance with the applicable noise requirements of part 
36 of this chapter and, in the case of an emissions change described in 
Sec. 21.93(c), show compliance with the applicable fuel venting and 
exhaust emissions requirements of part 34 of this chapter.
    (b) Each applicant for a supplemental type certificate must meet 
Sec. Sec. 21.33 and 21.53 with respect to each change in the type 
design.

[Amdt. 21-17, 32 FR 14927, Oct. 28, 1967, as amended by Amdt. 21-42, 40 
FR 1033, Jan. 6, 1975; Amdt. 21-52A, 45 FR 79009, Nov. 28, 1980; Amdt. 
21-61, 53 FR 3540, Feb. 5, 1988; Amdt. 21-68, 55 FR 32860, Aug. 10, 
1990; Amdt. 21-71, 57 FR 42854, Sept. 16, 1992; Amdt. 21-77, 65 FR 
36266, June 7, 2000]



Sec. 21.117  Issue of supplemental type certificates.

    (a) An applicant is entitled to a supplemental type certificate if 
the FAA finds that the applicant meets the requirements of Sec. Sec. 
21.113 and 21.115.
    (b) A supplemental type certificate consists of--
    (1) The approval by the FAA of a change in the type design of the 
product; and
    (2) The type certificate previously issued for the product.

[Doc. No. 5085, 29 FR 14568, Oct. 24, 1964, as amended by Amdt. 21-92, 
74 FR 53387, Oct. 16, 2009]



Sec. 21.119  Privileges.

    The holder of a supplemental type certificate may--
    (a) In the case of aircraft, obtain airworthiness certificates;
    (b) In the case of other products, obtain approval for installation 
on certificated aircraft; and
    (c) Obtain a production certificate in accordance with the 
requirements of subpart G of this part for the change in the type design 
approved by the supplemental type certificate.

[Doc. No. 5085, 29 FR 14568, Oct. 24, 1964, as amended by Amdt. 21-92, 
74 FR 53387, Oct. 16, 2009]



Sec. 21.120  Responsibility of supplemental type certificate holders
to provide written permission for alterations.

    A supplemental type certificate holder who allows a person to use 
the supplemental type certificate to alter an aircraft, aircraft engine, 
or propeller must provide that person with written permission acceptable 
to the FAA.

[Doc. No. FAA-2003-14825, 71 FR 52258, Sept. 1, 2006]

[[Page 150]]



               Subpart F_Production Under Type Certificate

    Source: Docket No. 5085, 29 FR 14568, Oct. 24, 1964, unless 
otherwise noted.



Sec. 21.121  Applicability.

    This subpart prescribes rules for production under a type 
certificate.



Sec. 21.122  Location of or change to manufacturing facilities.

    (a) A type certificate holder may utilize manufacturing facilities 
located outside of the United States if the FAA finds no undue burden in 
administering the applicable requirements of Title 49 U.S.C. and this 
subchapter.
    (b) The type certificate holder must obtain FAA approval before 
making any changes to the location of any of its manufacturing 
facilities.
    (c) The type certificate holder must immediately notify the FAA, in 
writing, of any change to the manufacturing facilities that may affect 
the inspection, conformity, or airworthiness of its product or article.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53387, Oct. 16, 2009; Amdt. 
21-92A, 75 FR 9095, Mar. 1, 2010]



Sec. 21.123  Production under type certificate.

    Each manufacturer of a product being manufactured under a type 
certificate must--
    (a) Maintain at the place of manufacture all information and data 
specified in Sec. Sec. 21.31 and 21.41;
    (b) Make each product and article thereof available for inspection 
by the FAA;
    (c) Maintain records of the completion of all inspections and tests 
required by Sec. Sec. 21.127, 21.128, and 21.129 for at least 5 years 
for the products and articles thereof manufactured under the approval 
and at least 10 years for critical components identified under Sec. 
45.15(c) of this chapter;
    (d) Allow the FAA to make any inspection or test, including any 
inspection or test at a supplier facility, necessary to determine 
compliance with this subchapter;
    (e) Mark the product in accordance with part 45 of this chapter, 
including any critical parts;
    (f) Identify any portion of that product (e.g., sub-assemblies, 
component parts, or replacement articles) that leave the manufacturer's 
facility as FAA approved with the manufacturer's part number and name, 
trademark, symbol, or other FAA-approved manufacturer's identification; 
and
    (g) Except as otherwise authorized by the FAA, obtain a production 
certificate for that product in accordance with subpart G of this part 
within 6 months after the date of issuance of the type certificate.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53387, Oct. 16, 2009]



Sec. 21.125  [Reserved]



Sec. 21.127  Tests: aircraft.

    (a) Each person manufacturing aircraft under a type certificate must 
establish an approved production flight test procedure and flight check-
off form, and in accordance with that form, flight test each aircraft 
produced.
    (b) Each production flight test procedure must include the 
following:
    (1) An operational check of the trim, controllability, or other 
flight characteristics to establish that the production aircraft has the 
same range and degree of control as the prototype aircraft.
    (2) An operational check of each part or system operated by the crew 
while in flight to establish that, during flight, instrument readings 
are within normal range.
    (3) A determination that all instruments are properly marked, and 
that all placards and required flight manuals are installed after flight 
test.
    (4) A check of the operational characteristics of the aircraft on 
the ground.
    (5) A check on any other items peculiar to the aircraft being tested 
that can best be done during the ground or flight operation of the 
aircraft.



Sec. 21.128  Tests: aircraft engines.

    (a) Each person manufacturing aircraft engines under a type 
certificate must subject each engine (except rocket engines for which 
the manufacturer

[[Page 151]]

must establish a sampling technique) to an acceptable test run that 
includes the following:
    (1) Break-in runs that include a determination of fuel and oil 
consumption and a determination of power characteristics at rated 
maximum continuous power or thrust and, if applicable, at rated takeoff 
power or thrust.
    (2) At least five hours of operation at rated maximum continuous 
power or thrust. For engines having a rated takeoff power or thrust 
higher than rated maximum continuous power or thrust, the five-hour run 
must include 30 minutes at rated takeoff power or thrust.
    (b) The test runs required by paragraph (a) of this section may be 
made with the engine appropriately mounted and using current types of 
power and thrust measuring equipment.

[Doc. No. 5085, 29 FR 14568, Oct. 24, 1964, as amended by Amdt. 21-5, 32 
FR 3735, Mar. 4, 1967]



Sec. 21.129  Tests: propellers.

    Each person manufacturing propellers under a type certificate must 
give each variable pitch propeller an acceptable functional test to 
determine if it operates properly throughout the normal range of 
operation.



Sec. 21.130  Statement of conformity.

    Each holder or licensee of a type certificate who manufactures a 
product under this subpart must provide, in a form and manner acceptable 
to the FAA, a statement that the product for which the type certificate 
has been issued conforms to its type certificate and is in a condition 
for safe operation.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53387, Oct. 16, 2009]



                    Subpart G_Production Certificates

    Source: Docket No. FAA-2006-25877, Amdt. 21-92, 74 FR 53387, Oct. 
16, 2009, unless otherwise noted.



Sec. 21.131  Applicability.

    This subpart prescribes--
    (a) Procedural requirements for issuing production certificates; and
    (b) Rules governing holders of those certificates.



Sec. 21.132  Eligibility.

    Any person may apply for a production certificate if that person 
holds, for the product concerned--
    (a) A current type certificate,
    (b) A supplemental type certificate, or
    (c) Rights to the benefits of that type certificate or supplemental 
type certificate under a licensing agreement.



Sec. 21.133  Application.

    Each applicant must apply for a production certificate in a form and 
manner prescribed by the FAA.



Sec. 21.135  Organization.

    Each applicant for or holder of a production certificate must 
provide the FAA with a document describing how its organization will 
ensure compliance with the provisions of this subpart. At a minimum, the 
document must describe assigned responsibilities and delegated 
authority, and the functional relationship of those responsible for 
quality to management and other organizational components.



Sec. 21.137  Quality system.

    Each applicant for or holder of a production certificate must 
establish and describe in writing a quality system that ensures that 
each product and article conforms to its approved design and is in a 
condition for safe operation. This quality system must include:
    (a) Design data control. Procedures for controlling design data and 
subsequent changes to ensure that only current, correct, and approved 
data is used.
    (b) Document control. Procedures for controlling quality system 
documents and data and subsequent changes to ensure that only current, 
correct, and approved documents and data are used.
    (c) Supplier control. Procedures that--
    (1) Ensure that each supplier-furnished product or article conforms 
to its approved design; and
    (2) Require each supplier to report to the production approval 
holder if a product or article has been released from that supplier and 
subsequently found not to conform to the applicable design data.

[[Page 152]]

    (d) Manufacturing process control. Procedures for controlling 
manufacturing processes to ensure that each product and article conforms 
to its approved design.
    (e) Inspecting and testing. Procedures for inspections and tests 
used to ensure that each product and article conforms to its approved 
design. These procedures must include the following, as applicable:
    (1) A flight test of each aircraft produced unless that aircraft 
will be exported as an unassembled aircraft.
    (2) A functional test of each aircraft engine and each propeller 
produced.
    (f) Inspection, measuring, and test equipment control. Procedures to 
ensure calibration and control of all inspection, measuring, and test 
equipment used in determining conformity of each product and article to 
its approved design. Each calibration standard must be traceable to a 
standard acceptable to the FAA.
    (g) Inspection and test status. Procedures for documenting the 
inspection and test status of products and articles supplied or 
manufactured to the approved design.
    (h) Nonconforming product and article control. (1) Procedures to 
ensure that only products or articles that conform to their approved 
design are installed on a type-certificated product. These procedures 
must provide for the identification, documentation, evaluation, 
segregation, and disposition of nonconforming products and articles. 
Only authorized individuals may make disposition determinations.
    (2) Procedures to ensure that discarded articles are rendered 
unusable.
    (i) Corrective and preventive actions. Procedures for implementing 
corrective and preventive actions to eliminate the causes of an actual 
or potential nonconformity to the approved design or noncompliance with 
the approved quality system.
    (j) Handling and storage. Procedures to prevent damage and 
deterioration of each product and article during handling, storage, 
preservation, and packaging.
    (k) Control of quality records. Procedures for identifying, storing, 
protecting, retrieving, and retaining quality records. A production 
approval holder must retain these records for at least 5 years for the 
products and articles manufactured under the approval and at least 10 
years for critical components identified under Sec. 45.15(c) of this 
chapter.
    (l) Internal audits. Procedures for planning, conducting, and 
documenting internal audits to ensure compliance with the approved 
quality system. The procedures must include reporting results of 
internal audits to the manager responsible for implementing corrective 
and preventive actions.
    (m) In-service feedback. Procedures for receiving and processing 
feedback on in-service failures, malfunctions, and defects. These 
procedures must include a process for assisting the design approval 
holder to--
    (1) Address any in-service problem involving design changes; and
    (2) Determine if any changes to the Instructions for Continued 
Airworthiness are necessary.
    (n) Quality escapes. Procedures for identifying, analyzing, and 
initiating appropriate corrective action for products or articles that 
have been released from the quality system and that do not conform to 
the applicable design data or quality system requirements.



Sec. 21.138  Quality manual.

    Each applicant for or holder of a production certificate must 
provide a manual describing its quality system to the FAA for approval. 
The manual must be in the English language and retrievable in a form 
acceptable to the FAA.



Sec. 21.139  Location of or change to manufacturing facilities.

    (a) An applicant may obtain a production certificate for 
manufacturing facilities located outside of the United States if the FAA 
finds no undue burden in administering the applicable requirements of 
Title 49 U.S.C. and this subchapter.
    (b) The production certificate holder must obtain FAA approval 
before making any changes to the location of any of its manufacturing 
facilities.
    (c) The production certificate holder must immediately notify the 
FAA, in

[[Page 153]]

writing, of any change to the manufacturing facilities that may affect 
the inspection, conformity, or airworthiness of its product or article.



Sec. 21.140  Inspections and tests.

    Each applicant for or holder of a production certificate must allow 
the FAA to inspect its quality system, facilities, technical data, and 
any manufactured products or articles and witness any tests, including 
any inspections or tests at a supplier facility, necessary to determine 
compliance with this subchapter.



Sec. 21.141  Issuance.

    The FAA issues a production certificate after finding that the 
applicant complies with the requirements of this subpart.



Sec. 21.142  Production limitation record.

    The FAA issues a production limitation record as part of a 
production certificate. The record lists the type certificate number and 
the model of every product that the production certificate holder is 
authorized to manufacture.



Sec. 21.143  Duration.

    A production certificate is effective until surrendered, suspended, 
revoked, or the FAA otherwise establishes a termination date.



Sec. 21.144  Transferability.

    The holder of a production certificate may not transfer the 
production certificate.



Sec. 21.145  Privileges.

    (a) The holder of a production certificate may--
    (1) Obtain an aircraft airworthiness certificate without further 
showing, except that the FAA may inspect the aircraft for conformity 
with the type design; or
    (2) In the case of other products, obtain approval from the FAA for 
installation on type-certificated aircraft.
    (b) Notwithstanding the provisions of Sec. 147.3 of this chapter, 
the holder of a production certificate for a primary category aircraft, 
or for a normal, utility, or acrobatic category aircraft of a type 
design that is eligible for a special airworthiness certificate in the 
primary category under Sec. 21.184(c), may--
    (1) Conduct training for persons in the performance of a special 
inspection and preventive maintenance program approved as a part of the 
aircraft's type design under Sec. 21.24(b), provided a person holding a 
mechanic certificate with appropriate airframe and powerplant ratings 
issued under part 65 of this chapter gives the training; and
    (2) Issue a certificate of competency to persons successfully 
completing the approved training program, provided the certificate 
specifies the aircraft make and model to which the certificate applies.



Sec. 21.146  Responsibility of holder.

    The holder of a production certificate must--
    (a) Amend the document required by Sec. 21.135 as necessary to 
reflect changes in the organization and provide these amendments to the 
FAA.
    (b) Maintain the quality system in compliance with the data and 
procedures approved for the production certificate;
    (c) Ensure that each completed product or article for which a 
production certificate has been issued, including primary category 
aircraft assembled under a production certificate by another person from 
a kit provided by the holder of the production certificate, presented 
for airworthiness certification or approval conforms to its approved 
design and is in a condition for safe operation;
    (d) Mark the product or article for which a certificate or approval 
has been issued. Marking must be in accordance with part 45 of this 
chapter, including any critical parts;
    (e) Identify any portion of the product or article (e.g., sub-
assemblies, component parts, or replacement articles) that leave the 
manufacturer's facility as FAA approved with the manufacturer's part 
number and name, trademark, symbol, or other FAA approved manufacturer's 
identification;
    (f) Have access to type design data necessary to determine 
conformity and airworthiness for each product and article produced under 
the production certificate;

[[Page 154]]

    (g) Retain its production certificate and make it available to the 
FAA upon request; and
    (h) Make available to the FAA information regarding all delegation 
of authority to suppliers.



Sec. 21.147  Amendment of production certificates.

    The holder of a production certificate must apply for an amendment 
to a production certificate in a form and manner prescribed by the FAA. 
The applicant for an amendment to a production certificate to add a type 
certificate or model, or both, must comply with the applicable 
requirements of Sec. Sec. 21.137, 21.138, and 21.150.



Sec. 21.150  Changes in quality system.

    After the issuance of a production certificate--
    (a) Each change to the quality system is subject to review by the 
FAA; and
    (b) The holder of a production certificate must immediately notify 
the FAA, in writing, of any change that may affect the inspection, 
conformity, or airworthiness of its product or article.



                  Subpart H_Airworthiness Certificates

    Source: Docket No. 5085, 29 FR 14569, Oct. 24, 1964, unless 
otherwise noted.



Sec. 21.171  Applicability.

    This subpart prescribes procedural requirements for the issue of 
airworthiness certificates.



Sec. 21.173  Eligibility.

    Any registered owner of a U.S.-registered aircraft (or the agent of 
the owner) may apply for an airworthiness certificate for that aircraft. 
An application for an airworthiness certificate must be made in a form 
and manner acceptable to the FAA, and may be submitted to any FAA 
office.

[Amdt. 21-26, 34 FR 15244, Sept. 30, 1969]



Sec. 21.175  Airworthiness certificates: classification.

    (a) Standard airworthiness certificates are airworthiness 
certificates issued for aircraft type certificated in the normal, 
utility, acrobatic, commuter, or transport category, and for manned free 
balloons, and for aircraft designated by the FAA as special classes of 
aircraft.
    (b) Special airworthiness certificates are primary, restricted, 
limited, light-sport, and provisional airworthiness certificates, 
special flight permits, and experimental certificates.

[Amdt. 21-21, 33 FR 6858, May 7, 1968, as amended by Amdt. 21-60, 52 FR 
8043, Mar. 13, 1987; Amdt. 21-70, 57 FR 41368, Sept. 9, 1992; Amdt. 21-
85, 69 FR 44861, July 27, 2004]



Sec. 21.177  Amendment or modification.

    An airworthiness certificate may be amended or modified only upon 
application to the FAA.



Sec. 21.179  Transferability.

    An airworthiness certificate is transferred with the aircraft.



Sec. 21.181  Duration.

    (a) Unless sooner surrendered, suspended, revoked, or a termination 
date is otherwise established by the FAA, airworthiness certificates are 
effective as follows:
    (1) Standard airworthiness certificates, special airworthiness 
certificates--primary category, and airworthiness certificates issued 
for restricted or limited category aircraft are effective as long as the 
maintenance, preventive maintenance, and alterations are performed in 
accordance with Parts 43 and 91 of this chapter and the aircraft are 
registered in the United States.
    (2) A special flight permit is effective for the period of time 
specified in the permit.
    (3) A special airworthiness certificate in the light-sport category 
is effective as long as--
    (i) The aircraft meets the definition of a light-sport aircraft;
    (ii) The aircraft conforms to its original configuration, except for 
those alterations performed in accordance with an applicable consensus 
standard and authorized by the aircraft's manufacturer or a person 
acceptable to the FAA;

[[Page 155]]

    (iii) The aircraft has no unsafe condition and is not likely to 
develop an unsafe condition; and
    (iv) The aircraft is registered in the United States.
    (4) An experimental certificate for research and development, 
showing compliance with regulations, crew training, or market surveys is 
effective for 1 year after the date of issue or renewal unless the FAA 
prescribes a shorter period. The duration of an experimental certificate 
issued for operating amateur-built aircraft, exhibition, air-racing, 
operating primary kit-built aircraft, or operating light-sport aircraft 
is unlimited, unless the FAA establishes a specific period for good 
cause.
    (b) The owner, operator, or bailee of the aircraft must, upon 
request, make it available for inspection by the FAA.
    (c) Upon suspension, revocation, or termination by order of the FAA 
of an airworthiness certificate, the owner, operator, or bailee of an 
aircraft must, upon request, surrender the certificate to the FAA.

[Amdt. 21-21, 33 FR 6858, May 7, 1968, as amended by Amdt. 21-49, 44 FR 
46781, Aug. 9, 1979; Amdt. 21-70, 57 FR 41368, Sept. 9, 1992; Amdt. 21-
85, 69 FR 44861, July 27, 2004]



Sec. 21.182  Aircraft identification.

    (a) Except as provided in paragraph (b) of this section, each 
applicant for an airworthiness certificate under this subpart must show 
that his aircraft is identified as prescribed in Sec. 45.11.
    (b) Paragraph (a) of this section does not apply to applicants for 
the following:
    (1) A special flight permit.
    (2) An experimental certificate for an aircraft not issued for the 
purpose of operating amateur-built aircraft, operating primary kit-built 
aircraft, or operating light-sport aircraft.
    (3) A change from one airworthiness classification to another, for 
an aircraft already identified as prescribed in Sec. 45.11.

[Amdt. 21-13, 32 FR 188, Jan. 10, 1967, as amended by Amdt. 21-51, 45 FR 
60170, Sept. 11, 1980; Amdt. 21-70, 57 FR 41368, Sept. 9, 1992; Amdt. 
21-85, 69 FR 44862, July 27, 2004]



Sec. 21.183  Issue of standard airworthiness certificates for normal,
utility, acrobatic, commuter, and transport category aircraft; manned
free balloons; and special classes of aircraft.

    (a) New aircraft manufactured under a production certificate. An 
applicant for a standard airworthiness certificate for a new aircraft 
manufactured under a production certificate is entitled to a standard 
airworthiness certificate without further showing, except that the FAA 
may inspect the aircraft to determine conformity to the type design and 
condition for safe operation.
    (b) New aircraft manufactured under type certificate. An applicant 
for a standard airworthiness certificate for a new aircraft manufactured 
under a type certificate is entitled to a standard airworthiness 
certificate upon presentation, by the holder or licensee of the type 
certificate, of the statement of conformity prescribed in Sec. 21.130 
if the FAA finds after inspection that the aircraft conforms to the type 
design and is in condition for safe operation.
    (c) Import aircraft. An applicant for a standard airworthiness 
certificate for an import aircraft is entitled to that certificate if--
    (1) The aircraft is type certificated in accordance with Sec. 21.21 
or Sec. 21.29 and produced under the authority of another State of 
Manufacture;
    (2) The State of Manufacture certifies, in accordance with the 
export provisions of an agreement with the United States for import of 
that aircraft, that the aircraft conforms to the type design and is in 
condition for safe operation; and
    (3) The FAA finds that the aircraft conforms to the type design and 
is in condition for safe operation.
    (d) Used aircraft and surplus aircraft of the U.S. Armed Forces. An 
applicant for a standard airworthiness certificate for a used aircraft 
or surplus aircraft of the U.S. Armed Forces is entitled to a standard 
airworthiness certificate if--
    (1) The applicant presents evidence to the FAA that the aircraft 
conforms to a type design approved under a type certificate or a 
supplemental type certificate and to applicable Airworthiness 
Directives;

[[Page 156]]

    (2) The aircraft (except an experimentally certificated aircraft 
that previously had been issued a different airworthiness certificate 
under this section) has been inspected in accordance with the 
performance rules for 100-hour inspections set forth in Sec. 43.15 of 
this chapter, or an equivalent performance standard acceptable to the 
FAA, and found airworthy by--
    (i) The manufacturer;
    (ii) The holder of a repair station certificate as provided in Part 
145 of this chapter;
    (iii) The holder of a mechanic certificate as authorized in Part 65 
of this chapter; or
    (iv) The holder of a certificate issued under part 121 of this 
chapter, and having a maintenance and inspection organization 
appropriate to the aircraft type; and
    (3) The FAA finds after inspection, that the aircraft conforms to 
the type design, and is in condition for safe operation.
    (e) Noise requirements. Notwithstanding all other provisions of this 
section, the following must be complied with for the original issuance 
of a standard airworthiness certificate:
    (1) For transport category large airplanes and jet (turbojet 
powered) airplanes that have not had any flight time before the dates 
specified in Sec. 36.1(d), no standard airworthiness certificate is 
originally issued under this section unless the FAA finds that the type 
design complies with the noise requirements in Sec. 36.1(d) in addition 
to the applicable airworthiness requirements in this section. For import 
airplanes, compliance with this paragraph is shown if the country in 
which the airplane was manufactured certifies, and the FAA finds, that 
Sec. 36.1(d) (or the applicable airplane noise requirements of the 
country in which the airplane was manufactured and any other 
requirements the FAA may prescribe to provide noise levels no greater 
than those provided by compliance with Sec. 36.1(d)) and paragraph (c) 
of this section are complied with.
    (2) For normal, utility, acrobatic, commuter, or transport category 
propeller driven small airplanes (except for those airplanes that are 
designed for ``agricultural aircraft operations'' (as defined in Sec. 
137.3 of this chapter, as effective on January 1, 1966) or for 
dispensing fire fighting materials to which Sec. 36.1583 of this 
chapter does not apply) that have not had any flight time before the 
applicable date specified in part 36 of this chapter, no standard 
airworthiness certificate is originally issued under this section unless 
the applicant shows that the type design complies with the applicable 
noise requirements of part 36 of this chapter in addition to the 
applicable airworthiness requirements in this section. For import 
airplanes, compliance with this paragraph is shown if the country in 
which the airplane was manufactured certifies, and the FAA finds, that 
the applicable requirements of part of this chapter (or the applicable 
airplane noise requirements of the country in which the airplane was 
manufactured and any other requirements the FAA may prescribe to provide 
noise levels no greater than those provided by compliance with the 
applicable requirements of part 36 of this chapter) and paragraph (c) of 
this section are complied with.
    (f) Passenger emergency exit requirements. Notwithstanding all other 
provisions of this section, each applicant for issuance of a standard 
airworthiness certificate for a transport category airplane manufactured 
after October 16, 1987, must show that the airplane meets the 
requirements of Sec. 25.807(c)(7) in effect on July 24, 1989. For the 
purposes of this paragraph, the date of manufacture of an airplane is 
the date the inspection acceptance records reflect that the airplane is 
complete and meets the FAA-approved type design data.
    (g) Fuel venting and exhaust emission requirements. Notwithstanding 
all other provisions of this section, and irrespective of the date of 
application, no airworthiness certificate is issued, on and after the 
dates specified in part 34 for the airplanes specified therein, unless 
the airplane complies with the applicable requirements of that part.
    (h) New aircraft manufactured under the provisions of Sec. 21.6(b). 
An applicant for a standard airworthiness certificate for a new aircraft 
manufactured under the provisions of Sec. 21.6(b) is entitled to

[[Page 157]]

a standard airworthiness certificate if--
    (1) The applicant presents evidence to the FAA that the aircraft 
conforms to a type design approved under a type certificate or 
supplemental type certificate and to applicable Airworthiness 
Directives;
    (2) The aircraft has been inspected in accordance with the 
performance rules for a 100-hour inspections set forth in Sec. 43.15 of 
this chapter and found airworthy by a person specified in paragraph 
(d)(2) of this section; and
    (3) The FAA finds after inspection, that the aircraft conforms to 
the type design, and is in condition for safe operation.

[Amdt. 21-17, 32 FR 14927, Oct. 28, 1967]

    Editorial Note: For Federal Register citations affecting Sec. 
21.183, see the List of CFR Sections Affected, which appears in the 
Finding Aids section of the printed volume and at www.fdsys.gov.



Sec. 21.184  Issue of special airworthiness certificates for primary
category aircraft.

    (a) New primary category aircraft manufactured under a production 
certificate. An applicant for an original, special airworthiness 
certificate-primary category for a new aircraft that meets the criteria 
of Sec. 21.24(a)(1), manufactured under a production certificate, 
including aircraft assembled by another person from a kit provided by 
the holder of the production certificate and under the supervision and 
quality control of that holder, is entitled to a special airworthiness 
certificate without further showing, except that the FAA may inspect the 
aircraft to determine conformity to the type design and condition for 
safe operation.
    (b) Imported aircraft. An applicant for a special airworthiness 
certificate-primary category for an imported aircraft type certificated 
under Sec. 21.29 is entitled to a special airworthiness certificate if 
the civil airworthiness authority of the country in which the aircraft 
was manufactured certifies, and the FAA finds after inspection, that the 
aircraft conforms to an approved type design that meets the criteria of 
Sec. 21.24(a)(1) and is in a condition for safe operation.
    (c) Aircraft having a current standard airworthiness certificate. An 
applicant for a special airworthiness certificate-primary category, for 
an aircraft having a current standard airworthiness certificate that 
meets the criteria of Sec. 21.24(a)(1), may obtain the primary category 
certificate in exchange for its standard airworthiness certificate 
through the supplemental type certification process. For the purposes of 
this paragraph, a current standard airworthiness certificate means that 
the aircraft conforms to its approved normal, utility, or acrobatic type 
design, complies with all applicable airworthiness directives, has been 
inspected and found airworthy within the last 12 calendar months in 
accordance with Sec. 91.409(a)(1) of this chapter, and is found to be 
in a condition for safe operation by the FAA.
    (d) Other aircraft. An applicant for a special airworthiness 
certificate-primary category for an aircraft that meets the criteria of 
Sec. 21.24(a)(1), and is not covered by paragraph (a), (b), or (c) of 
this section, is entitled to a special airworthiness certificate if--
    (1) The applicant presents evidence to the FAA that the aircraft 
conforms to an approved primary, normal, utility, or acrobatic type 
design, including compliance with all applicable airworthiness 
directives;
    (2) The aircraft has been inspected and found airworthy within the 
past 12 calendar months in accordance with Sec. 91.409(a)(1) of this 
chapter and;
    (3) The aircraft is found by the FAA to conform to an approved type 
design and to be in a condition for safe operation.
    (e) Multiple-category airworthiness certificates in the primary 
category and any other category will not be issued; a primary category 
aircraft may hold only one airworthiness certificate.

[Doc. No. 23345, 57 FR 41368, Sept. 9, 1992, as amended by Amdt. 21-70, 
57 FR 43776, Sept. 22, 1992]



Sec. 21.185  Issue of airworthiness certificates for restricted
category aircraft.

    (a) Aircraft manufactured under a production certificate or type 
certificate. An

[[Page 158]]

applicant for the original issue of a restricted category airworthiness 
certificate for an aircraft type certificated in the restricted 
category, that was not previously type certificated in any other 
category, must comply with the appropriate provisions of Sec. 21.183.
    (b) Other aircraft. An applicant for a restricted category 
airworthiness certificate for an aircraft type certificated in the 
restricted category, that was either a surplus aircraft of the Armed 
Forces or previously type certificated in another category, is entitled 
to an airworthiness certificate if the aircraft has been inspected by 
the FAA and found by him to be in a good state of preservation and 
repair and in a condition for safe operation.
    (c) Import aircraft. An applicant for the original issue of a 
special airworthiness certificate for a restricted category import 
aircraft is entitled to that certificate if--
    (1) The aircraft is type-certificated in accordance with Sec. 21.25 
or Sec. 21.29 and produced under the authority of another State of 
Manufacture;
    (2) The State of Manufacture certifies, in accordance with the 
export provisions of an agreement with the United States for import of 
that aircraft that the aircraft conforms to the type design and is in 
condition for safe operation; and
    (3) The FAA finds that the aircraft conforms to the type design and 
is in condition for safe operation.
    (d) Noise requirements. For propeller-driven small airplanes (except 
airplanes designed for ``agricultural aircraft operations,'' as defined 
in Sec. 137.3 of this chapter, as effective on January 1, 1966, or for 
dispensing fire fighting materials) that have not had any flight time 
before the applicable date specified in Part 36 of this chapter, and 
notwithstanding the other provisions of this section, no original 
restricted category airworthiness certificate is issued under this 
section unless the FAA finds that the type design complies with the 
applicable noise requirements of Part 36 of this chapter in addition to 
the applicable airworthiness requirements of this section. For import 
airplanes, compliance with this paragraph is shown if the country in 
which the airplane was manufactured certifies, and the FAA finds, that 
the applicable requirements of Part 36 of this chapter (or the 
applicable airplane noise requirements of the country in which the 
airplane was manufactured and any other requirements the FAA may 
prescribe to provide noise levels no greater than those provided by 
compliance with the applicable requirements of Part 36 of this chapter) 
and paragraph (c) of this section are complied with.

[Amdt. 21-10, 31 FR 9211, July 6, 1966; as amended by Amdt. 21-32, 35 FR 
10202, June 23, 1970; Amdt. 21-42, 40 FR 1034, Jan. 6, 1975; Amdt. 21-
92, 74 FR 53389, Oct. 16, 2009; Amdt. 21-92, 74 FR 53389, Oct. 16, 2009; 
Amdt. 21-92A, 75 FR 9095, Mar. 1, 2010]



Sec. 21.187  Issue of multiple airworthiness certification.

    (a) An applicant for an airworthiness certificate in the restricted 
category, and in one or more other categories except primary category, 
is entitled to the certificate, if--
    (1) He shows compliance with the requirements for each category, 
when the aircraft is in the configuration for that category; and
    (2) He shows that the aircraft can be converted from one category to 
another by removing or adding equipment by simple mechanical means.
    (b) The operator of an aircraft certificated under this section must 
have the aircraft inspected by the FAA, or by a certificated mechanic 
with an appropriate airframe rating, to determine airworthiness each 
time the aircraft is converted from the restricted category to another 
category for the carriage of passengers for compensation or hire, unless 
the FAA finds this unnecessary for safety in a particular case.
    (c) The aircraft complies with the applicable requirements of part 
34.

[Doc. No. 5085, 29 FR 14569, Oct. 24, 1964, as amended by Amdt. 21-68, 
55 FR 32860, Aug. 10, 1990; Amdt. 21-70, 57 FR 41369, Sept. 9, 1992]



Sec. 21.189  Issue of airworthiness certificate for limited category
aircraft.

    (a) An applicant for an airworthiness certificate for an aircraft in 
the limited category is entitled to the certificate when--

[[Page 159]]

    (1) He shows that the aircraft has been previously issued a limited 
category type certificate and that the aircraft conforms to that type 
certificate; and
    (2) The FAA finds, after inspection (including a flight check by the 
applicant), that the aircraft is in a good state of preservation and 
repair and is in a condition for safe operation.
    (b) The FAA prescribes limitations and conditions necessary for safe 
operation.

[Doc. No. 5085, 29 FR 14570, Oct. 24, 1964, as amended by Amdt. 21-4, 30 
FR 9437, July 29, 1965]



Sec. 21.190  Issue of a special airworthiness certificate for a 
light-sport category aircraft.

    (a) Purpose. The FAA issues a special airworthiness certificate in 
the light-sport category to operate a light-sport aircraft, other than a 
gyroplane.
    (b) Eligibility. To be eligible for a special airworthiness 
certificate in the light-sport category:
    (1) An applicant must provide the FAA with--
    (i) The aircraft's operating instructions;
    (ii) The aircraft's maintenance and inspection procedures;
    (iii) The manufacturer's statement of compliance as described in 
paragraph (c) of this section; and
    (iv) The aircraft's flight training supplement.
    (2) The aircraft must not have been previously issued a standard, 
primary, restricted, limited, or provisional airworthiness certificate, 
or an equivalent airworthiness certificate issued by a foreign civil 
aviation authority.
    (3) The aircraft must be inspected by the FAA and found to be in a 
condition for safe operation.
    (c) Manufacturer's statement of compliance for light-sport category 
aircraft. The manufacturer's statement of compliance required in 
paragraph (b)(1)(iii) of this section must--
    (1) Identify the aircraft by make and model, serial number, class, 
date of manufacture, and consensus standard used;
    (2) State that the aircraft meets the provisions of the identified 
consensus standard;
    (3) State that the aircraft conforms to the manufacturer's design 
data, using the manufacturer's quality assurance system that meets the 
identified consensus standard;
    (4) State that the manufacturer will make available to any 
interested person the following documents that meet the identified 
consensus standard:
    (i) The aircraft's operating instructions.
    (ii) The aircraft's maintenance and inspection procedures.
    (iii) The aircraft's flight training supplement.
    (5) State that the manufacturer will monitor and correct safety-of-
flight issues through the issuance of safety directives and a continued 
airworthiness system that meets the identified consensus standard;
    (6) State that at the request of the FAA, the manufacturer will 
provide unrestricted access to its facilities; and
    (7) State that the manufacturer, in accordance with a production 
acceptance test procedure that meets an applicable consensus standard 
has--
    (i) Ground and flight tested the aircraft;
    (ii) Found the aircraft performance acceptable; and
    (iii) Determined that the aircraft is in a condition for safe 
operation.
    (d) Light-sport aircraft manufactured outside the United States. For 
aircraft manufactured outside of the United States to be eligible for a 
special airworthiness certificate in the light-sport category, an 
applicant must meet the requirements of paragraph (b) of this section 
and provide to the FAA evidence that--
    (1) The aircraft was manufactured in a country with which the United 
States has a Bilateral Airworthiness Agreement concerning airplanes or 
Bilateral Aviation Safety Agreement with associated Implementation 
Procedures for Airworthiness concerning airplanes, or an equivalent 
airworthiness agreement; and
    (2) The aircraft is eligible for an airworthiness certificate, 
flight authorization, or other similar certification in its country of 
manufacture.

[Amdt. 21-85, 69 FR 44862, July 27, 2004]

[[Page 160]]



Sec. 21.191  Experimental certificates.

    Experimental certificates are issued for the following purposes:
    (a) Research and development. Testing new aircraft design concepts, 
new aircraft equipment, new aircraft installations, new aircraft 
operating techniques, or new uses for aircraft.
    (b) Showing compliance with regulations. Conducting flight tests and 
other operations to show compliance with the airworthiness regulations 
including flights to show compliance for issuance of type and 
supplemental type certificates, flights to substantiate major design 
changes, and flights to show compliance with the function and 
reliability requirements of the regulations.
    (c) Crew training. Training of the applicant's flight crews.
    (d) Exhibition. Exhibiting the aircraft's flight capabilities, 
performance, or unusual characteristics at air shows, motion picture, 
television, and similar productions, and the maintenance of exhibition 
flight proficiency, including (for persons exhibiting aircraft) flying 
to and from such air shows and productions.
    (e) Air racing. Participating in air races, including (for such 
participants) practicing for such air races and flying to and from 
racing events.
    (f) Market surveys. Use of aircraft for purposes of conducting 
market surveys, sales demonstrations, and customer crew training only as 
provided in Sec. 21.195.
    (g) Operating amateur-built aircraft. Operating an aircraft the 
major portion of which has been fabricated and assembled by persons who 
undertook the construction project solely for their own education or 
recreation.
    (h) Operating primary kit-built aircraft. Operating a primary 
category aircraft that meets the criteria of Sec. 21.24(a)(1) that was 
assembled by a person from a kit manufactured by the holder of a 
production certificate for that kit, without the supervision and quality 
control of the production certificate holder under Sec. 21.184(a).
    (i) Operating light-sport aircraft. Operating a light-sport aircraft 
that--
    (1) Has not been issued a U.S. or foreign airworthiness certificate 
and does not meet the provisions of Sec. 103.1 of this chapter. An 
experimental certificate will not be issued under this paragraph for 
these aircraft after January 31, 2008;
    (2) Has been assembled--
    (i) From an aircraft kit for which the applicant can provide the 
information required by Sec. 21.193(e); and
    (ii) In accordance with manufacturer's assembly instructions that 
meet an applicable consensus standard; or
    (3) Has been previously issued a special airworthiness certificate 
in the light-sport category under Sec. 21.190.

[Amdt. 21-21, 38 FR 6858, May 7, 1968, as amended by Amdt. 21-57, 49 FR 
39651, Oct. 9, 1984; Amdt. 21-70, 57 FR 41369, Sept. 9, 1992; Amdt. 21-
85, 69 FR 44862, July 27, 2004; Amdt. 21-85, 69 FR 53336, Sept. 1, 2004]



Sec. 21.193  Experimental certificates: general.

    An applicant for an experimental certificate must submit the 
following information:
    (a) A statement, in a form and manner prescribed by the FAA setting 
forth the purpose for which the aircraft is to be used.
    (b) Enough data (such as photographs) to identify the aircraft.
    (c) Upon inspection of the aircraft, any pertinent information found 
necessary by the FAA to safeguard the general public.
    (d) In the case of an aircraft to be used for experimental 
purposes--
    (1) The purpose of the experiment;
    (2) The estimated time or number of flights required for the 
experiment;
    (3) The areas over which the experiment will be conducted; and
    (4) Except for aircraft converted from a previously certificated 
type without appreciable change in the external configuration, three-
view drawings or three-view dimensioned photographs of the aircraft.
    (e) In the case of a light-sport aircraft assembled from a kit to be 
certificated in accordance with Sec. 21.191(i)(2), an applicant must 
provide the following:
    (1) Evidence that an aircraft of the same make and model was 
manufactured and assembled by the aircraft kit manufacturer and issued a 
special airworthiness certificate in the light-sport category.

[[Page 161]]

    (2) The aircraft's operating instructions.
    (3) The aircraft's maintenance and inspection procedures.
    (4) The manufacturer's statement of compliance for the aircraft kit 
used in the aircraft assembly that meets Sec. 21.190(c), except that 
instead of meeting Sec. 21.190(c)(7), the statement must identify 
assembly instructions for the aircraft that meet an applicable consensus 
standard.
    (5) The aircraft's flight training supplement.
    (6) In addition to paragraphs (e)(1) through (e)(5) of this section, 
for an aircraft kit manufactured outside of the United States, evidence 
that the aircraft kit was manufactured in a country with which the 
United States has a Bilateral Airworthiness Agreement concerning 
airplanes or a Bilateral Aviation Safety Agreement with associated 
Implementation Procedures for Airworthiness concerning airplanes, or an 
equivalent airworthiness agreement.

[Doc. No. 5085, 29 FR 14569, Oct. 24, 1964, as amended by Amdt. 21-85, 
69 FR 44862, July 27, 2004]



Sec. 21.195  Experimental certificates: Aircraft to be used for market
surveys, sales demonstrations, and customer crew training.

    (a) A manufacturer of aircraft manufactured within the United States 
may apply for an experimental certificate for an aircraft that is to be 
used for market surveys, sales demonstrations, or customer crew 
training.
    (b) A manufacturer of aircraft engines who has altered a type 
certificated aircraft by installing different engines, manufactured by 
him within the United States, may apply for an experimental certificate 
for that aircraft to be used for market surveys, sales demonstrations, 
or customer crew training, if the basic aircraft, before alteration, was 
type certificated in the normal, acrobatic, commuter, or transport 
category.
    (c) A person who has altered the design of a type certificated 
aircraft may apply for an experimental certificate for the altered 
aircraft to be used for market surveys, sales demonstrations, or 
customer crew training if the basic aircraft, before alteration, was 
type certificated in the normal, utility, acrobatic, or transport 
category.
    (d) An applicant for an experimental certificate under this section 
is entitled to that certificate if, in addition to meeting the 
requirements of Sec. 21.193--
    (1) He has established an inspection and maintenance program for the 
continued airworthiness of the aircraft; and
    (2) The applicant shows that the aircraft has been flown for at 
least 50 hours, or for at least 5 hours if it is a type certificated 
aircraft which has been modified. The FAA may reduce these operational 
requirements if the applicant provides adequate justification.

[Amdt. 21-21, 33 FR 6858, May 7, 1968, as amended by Amdt. 21-28, 35 FR 
2818, Feb. 11, 1970; Amdt. 21-57, 49 FR 39651, Oct. 9, 1984; Amdt. 21-
59, 52 FR 1836, Jan. 15, 1987; Amdt. 21-92, 74 FR 53389, Oct. 16, 2009]



Sec. 21.197  Special flight permits.

    (a) A special flight permit may be issued for an aircraft that may 
not currently meet applicable airworthiness requirements but is capable 
of safe flight, for the following purposes:
    (1) Flying the aircraft to a base where repairs, alterations, or 
maintenance are to be performed, or to a point of storage.
    (2) Delivering or exporting the aircraft.
    (3) Production flight testing new production aircraft.
    (4) Evacuating aircraft from areas of impending danger.
    (5) Conducting customer demonstration flights in new production 
aircraft that have satisfactorily completed production flight tests.
    (b) A special flight permit may also be issued to authorize the 
operation of an aircraft at a weight in excess of its maximum 
certificated takeoff weight for flight beyond the normal range over 
water, or over land areas where adequate landing facilities or 
appropriate fuel is not available. The excess weight that may be 
authorized under this paragraph is limited to the additional fuel, fuel-
carrying facilities, and navigation equipment necessary for the flight.

[[Page 162]]

    (c) Upon application, as prescribed in Sec. Sec. 91.1017 or 119.51 
of this chapter, a special flight permit with a continuing authorization 
may be issued for aircraft that may not meet applicable airworthiness 
requirements, but are capable of safe flight for the purpose of flying 
aircraft to a base where maintenance or alterations are to be performed. 
The permit issued under this paragraph is an authorization, including 
conditions and limitations for flight, which is set forth in the 
certificate holder's operations specifications. The permit issued under 
this paragraph may be issued to--
    (1) Certificate holders authorized to conduct operations under part 
119 of this chapter, that have an approved program for continuing flight 
authorization; or
    (2) Management specification holders authorized to conduct 
operations under part 91, subpart K of this chapter for those aircraft 
they operate and maintain under a continuous airworthiness maintenance 
program prescribed by Sec. 91.1411 of this chapter.

[Doc. No. 5085, 29 FR 14570, Oct. 24, 1964, as amended by Amdt. 21-21, 
33 FR 6859, May 7, 1968; Amdt. 21-51, 45 FR 60170, Sept. 11, 1980; Amdt. 
21-54, 46 FR 37878, July 23, 1981; Amdt. 21-79, 66 FR 21066, Apr. 27, 
2001; Amdt. 21-84, 68 FR 54559, Sept. 17, 2003; Amdt. 21-87, 71 FR 536, 
Jan. 4, 2006; Amdt. 21-92, 74 FR 53389, Oct. 16, 2009]



Sec. 21.199  Issue of special flight permits.

    (a) Except as provided in Sec. 21.197(c), an applicant for a 
special flight permit must submit a statement in a form and manner 
prescribed by the FAA, indicating--
    (1) The purpose of the flight.
    (2) The proposed itinerary.
    (3) The crew required to operate the aircraft and its equipment, 
e.g., pilot, co-pilot, navigator, etc.
    (4) The ways, if any, in which the aircraft does not comply with the 
applicable airworthiness requirements.
    (5) Any restriction the applicant considers necessary for safe 
operation of the aircraft.
    (6) Any other information considered necessary by the FAA for the 
purpose of prescribing operating limitations.
    (b) The FAA may make, or require the applicant to make appropriate 
inspections or tests necessary for safety.

[Doc. No. 5085, 29 FR 14570, Oct. 24, 1964, as amended by Amdt. 21-21, 
33 FR 6859, May 7, 1968; Amdt. 21-22, 33 FR 11901, Aug. 22, 1968]



            Subpart I_Provisional Airworthiness Certificates

    Source: Docket No. 5085, 29 FR 14571, Oct. 24, 1964, unless 
otherwise noted.



Sec. 21.211  Applicability.

    This subpart prescribes procedural requirements for the issue of 
provisional airworthiness certificates.



Sec. 21.213  Eligibility.

    (a) A manufacturer who is a United States citizen may apply for a 
Class I or Class II provisional airworthiness certificate for aircraft 
manufactured by him within the U.S.
    (b) Any holder of an air carrier operating certificate under Part 
121 of this chapter who is a United States citizen may apply for a Class 
II provisional airworthiness certificate for transport category aircraft 
that meet either of the following:
    (1) The aircraft has a current Class II provisional type certificate 
or an amendment thereto.
    (2) The aircraft has a current provisional amendment to a type 
certificate that was preceded by a corresponding Class II provisional 
type certificate.
    (c) An aircraft engine manufacturer who is a United States citizen 
and who has altered a type certificated aircraft by installing different 
type certificated engines, manufactured by him within the United States, 
may apply for a Class I provisional airworthiness certificate for that 
aircraft, if the basic aircraft, before alteration, was type 
certificated in the normal, utility, acrobatic, commuter, or transport 
category.

[Doc. No. 5085, 29 FR 14571, Oct. 24, 1964, as amended by Amdt. 21-59, 
52 FR 1836, Jan. 15, 1987; Amdt. 21-79, 66 FR 21066, Apr. 27, 2001]

[[Page 163]]



Sec. 21.215  Application.

    Applications for provisional airworthiness certificates must be 
submitted to the Manufacturing Inspection District Office in the 
geographic area in which the manufacturer or air carrier is located. The 
application must be accompanied by the pertinent information specified 
in this subpart.

[Amdt. 21-67, 54 FR 39291, Sept. 25, 1989; 54 FR 52872, Dec. 22, 1989]



Sec. 21.217  Duration.

    Unless sooner surrendered, superseded, revoked, or otherwise 
terminated, provisional airworthiness certificates are effective for the 
duration of the corresponding provisional type certificate, amendment to 
a provisional type certificate, or provisional amendment to the type 
certificate.



Sec. 21.219  Transferability.

    Class I provisional airworthiness certificates are not transferable. 
Class II provisional airworthiness certificates may be transferred to an 
air carrier eligible to apply for a certificate under Sec. 21.213(b).



Sec. 21.221  Class I provisional airworthiness certificates.

    (a) Except as provided in Sec. 21.225, an applicant is entitled to 
a Class I provisional airworthiness certificate for an aircraft for 
which a Class I provisional type certificate has been issued if--
    (1) He meets the eligibility requirements of Sec. 21.213 and he 
complies with this section; and
    (2) The FAA finds that there is no feature, characteristic or 
condition of the aircraft that would make the aircraft unsafe when 
operated in accordance with the limitations established in Sec. Sec. 
21.81(e) and 91.317 of this subchapter.
    (b) The manufacturer must hold a provisional type certificate for 
the aircraft.
    (c) The manufacturer must submit a statement that the aircraft 
conforms to the type design corresponding to the provisional type 
certificate and has been found by him to be in safe operating condition 
under all applicable limitations.
    (d) The aircraft must be flown at least five hours by the 
manufacturer.
    (e) The aircraft must be supplied with a provisional aircraft flight 
manual or other document and appropriate placards containing the 
limitations established by Sec. Sec. 21.81(e) and 91.317.

[Doc. No. 5085, 29 FR 14571, Oct. 24, 1964, as amended by Amdt. 21-66, 
54 FR 34329, Aug. 18, 1989]



Sec. 21.223  Class II provisional airworthiness certificates.

    (a) Except as provided in Sec. 21.225, an applicant is entitled to 
a Class II provisional airworthiness certificate for an aircraft for 
which a Class II provisional type certificate has been issued if--
    (1) He meets the eligibility requirements of Sec. 21.213 and he 
complies with this section; and
    (2) The FAA finds that there is no feature, characteristic, or 
condition of the aircraft that would make the aircraft unsafe when 
operated in accordance with the limitations established in Sec. Sec. 
21.83(h), 91.317, and 121.207 of this chapter.
    (b) The applicant must show that a Class II provisional type 
certificate for the aircraft has been issued to the manufacturer.
    (c) The applicant must submit a statement by the manufacturer that 
the aircraft has been manufactured under a quality system adequate to 
ensure that the aircraft conforms to the type design corresponding with 
the provisional type certificate.
    (d) The applicant must submit a statement that the aircraft has been 
found by him to be in a safe operating condition under the applicable 
limitations.
    (e) The aircraft must be flown at least five hours by the 
manufacturer.
    (f) The aircraft must be supplied with a provisional aircraft flight 
manual containing the limitations established by Sec. Sec. 21.83(h), 
91.317, and 121.207 of this chapter.

[Doc. No. 5085, 29 FR 14571, Oct. 24, 1964, as amended by Amdt. 21-12, 
31 FR 13389, Oct. 15, 1966; Amdt. 21-66, 54 FR 34329, Aug. 18, 1989; 
Amdt. 21-92, 74 FR 53390, Oct. 16, 2009]

[[Page 164]]



Sec. 21.225  Provisional airworthiness certificates corresponding with
provisional amendments to type certificates.

    (a) An applicant is entitled to a Class I or a Class II provisional 
airworthiness certificate, for an aircraft, for which a provisional 
amendment to the type certificate has been issued, if--
    (1) He meets the eligibility requirements of Sec. 21.213 and he 
complies with this section; and
    (2) The FAA finds that there is no feature, characteristic, or 
condition of the aircraft, as modified in accordance with the 
provisionally amended type certificate, that would make the aircraft 
unsafe when operated in accordance with the applicable limitations 
established in Sec. Sec. 21.85(g), 91.317, and 121.207 of this chapter.
    (b) The applicant must show that the modification was made under a 
quality system adequate to ensure that the modification conforms to the 
provisionally amended type certificate.
    (c) The applicant must submit a statement that the aircraft has been 
found by him to be in a safe operating condition under the applicable 
limitations.
    (d) The aircraft must be flown at least five hours by the 
manufacturer.
    (e) The aircraft must be supplied with a provisional aircraft flight 
manual or other document and appropriate placards containing the 
limitations required by Sec. Sec. 21.85(g), 91.317, and 121.207 of this 
chapter.

[Doc. No. 5085, 29 FR 14571, Oct. 24, 1964, as amended by Amdt. 21-12, 
31 FR 13389, Oct. 15, 1966; Amdt. 21-66, 54 FR 34329, Aug. 18, 1989; 
Amdt. 21-92, 74 FR 53390, Oct. 16, 2009]

Subpart J [Reserved]



                 Subpart K_Parts Manufacturer Approvals

    Source: Docket No. FAA-2006-25877, Amdt. 21-92, 74 FR 53390, Oct. 
16, 2009, unless otherwise noted.



Sec. 21.301  Applicability.

    This subpart prescribes--
    (a) Procedural requirements for issuing PMAs; and
    (b) Rules governing holders of PMAs.



Sec. 21.303  Application.

    (a) The applicant for a PMA must apply in a form and manner 
prescribed by the FAA, and include the following:
    (1) The identity of the product on which the article is to be 
installed.
    (2) The name and address of the manufacturing facilities at which 
these articles are to be manufactured.
    (3) The design of the article, which consists of--
    (i) Drawings and specifications necessary to show the configuration 
of the article; and
    (ii) Information on dimensions, materials, and processes necessary 
to define the structural strength of the article.
    (4) Test reports and computations necessary to show that the design 
of the article meets the airworthiness requirements of this subchapter. 
The test reports and computations must be applicable to the product on 
which the article is to be installed, unless the applicant shows that 
the design of the article is identical to the design of a article that 
is covered under a type certificate. If the design of the article was 
obtained by a licensing agreement, the applicant must provide evidence 
of that agreement.
    (5) An applicant for a PMA based on test reports and computations 
must provide a statement certifying that the applicant has complied with 
the airworthiness requirements of this subchapter.
    (b) Each applicant for a PMA must make all inspections and tests 
necessary to determine--
    (1) Compliance with the applicable airworthiness requirements;
    (2) That materials conform to the specifications in the design;
    (3) That the article conforms to its approved design; and
    (4) That the manufacturing processes, construction, and assembly 
conform to those specified in the design.



Sec. 21.305  Organization.

    Each applicant for or holder of a PMA must provide the FAA with a 
document describing how its organization will ensure compliance with the 
provisions of this subpart. At a minimum, the document must describe 
assigned responsibilities and delegated authority, and the functional 
relationship of

[[Page 165]]

those responsible for quality to management and other organizational 
components.



Sec. 21.307  Quality system.

    Each applicant for or holder of a PMA must establish a quality 
system that meets the requirements of Sec. 21.137.



Sec. 21.308  Quality manual.

    Each applicant for or holder of a PMA must provide a manual 
describing its quality system to the FAA for approval. The manual must 
be in the English language and retrievable in a form acceptable to the 
FAA.



Sec. 21.309  Location of or change to manufacturing facilities.

    (a) An applicant may obtain a PMA for manufacturing facilities 
located outside of the United States if the FAA finds no undue burden in 
administering the applicable requirements of Title 49 U.S.C. and this 
subchapter.
    (b) The PMA holder must obtain FAA approval before making any 
changes to the location of any of its manufacturing facilities.
    (c) The PMA holder must immediately notify the FAA, in writing, of 
any change to the manufacturing facilities that may affect the 
inspection, conformity, or airworthiness of its PMA article.



Sec. 21.310  Inspections and tests.

    (a) Each applicant for or holder of a PMA must allow the FAA to 
inspect its quality system, facilities, technical data, and any 
manufactured articles and witness any tests, including any inspections 
or tests at a supplier facility, necessary to determine compliance with 
this subchapter.
    (b) Unless otherwise authorized by the FAA, the applicant or 
holder--
    (1) May not present any article to the FAA for an inspection or test 
unless compliance with Sec. 21.303(b)(2) through (4) has been shown for 
that article; and
    (2) May not make any change to an article between the time that 
compliance with Sec. 21.303(b)(2) through (4) is shown for that article 
and the time that the article is presented to the FAA for the inspection 
or test.



Sec. 21.311  Issuance.

    The FAA issues a PMA after finding that the applicant complies with 
the requirements of this subpart and the design complies with the 
requirements of this chapter applicable to the product on which the 
article is to be installed.



Sec. 21.313  Duration.

    A PMA is effective until surrendered, withdrawn, or the FAA 
otherwise terminates it.



Sec. 21.314  Transferability.

    The holder of a PMA may not transfer the PMA.



Sec. 21.316  Responsibility of holder.

    Each holder of a PMA must--
    (a) Amend the document required by Sec. 21.305 as necessary to 
reflect changes in the organization and provide these amendments to the 
FAA;
    (b) Maintain the quality system in compliance with the data and 
procedures approved for the PMA;
    (c) Ensure that each PMA article conforms to its approved design and 
is in a condition for safe operation;
    (d) Mark the PMA article for which an approval has been issued. 
Marking must be in accordance with part 45 of this chapter, including 
any critical parts;
    (e) Identify any portion of the PMA article (e.g., sub-assemblies, 
component parts, or replacement articles) that leave the manufacturer's 
facility as FAA approved with the manufacturer's part number and name, 
trademark, symbol, or other FAA approved manufacturer's identification;
    (f) Have access to design data necessary to determine conformity and 
airworthiness for each article produced under the PMA;
    (g) Retain each document granting PMA and make it available to the 
FAA upon request; and
    (h) Make available to the FAA information regarding all delegation 
of authority to suppliers.



Sec. 21.319  Design changes.

    (a) Classification of design changes. (1) A ``minor change'' to the 
design of an article produced under a PMA is one

[[Page 166]]

that has no appreciable effect on the approval basis.
    (2) A ``major change'' to the design of an article produced under a 
PMA is any change that is not minor.
    (b) Approval of design changes. (1) Minor changes to the basic 
design of a PMA may be approved using a method acceptable to the FAA.
    (2) The PMA holder must obtain FAA approval of any major change 
before including it in the design of an article produced under a PMA.



Sec. 21.320  Changes in quality system.

    After the issuance of a PMA--
    (a) Each change to the quality system is subject to review by the 
FAA; and
    (b) The holder of the PMA must immediately notify the FAA, in 
writing, of any change that may affect the inspection, conformity, or 
airworthiness of its article.



                Subpart L_Export Airworthiness Approvals

    Source: Docket No. FAA-2006-25877, Amdt. 21-92, 74 FR 53391, Oct. 
16, 2009, unless otherwise noted.



Sec. 21.321  Applicability.

    This subpart prescribes--
    (a) Procedural requirements for issuing export airworthiness 
approvals; and
    (b) Rules governing the holders of those approvals.



Sec. 21.325  Export airworthiness approvals.

    (a) An export airworthiness approval for an aircraft is issued in 
the form of an export certificate of airworthiness. This certificate 
does not authorize operation of that aircraft.
    (b) The FAA prescribes the form and manner in which an export 
airworthiness approval for an aircraft engine, propeller, or article is 
issued.
    (c) If the FAA finds no undue burden in administering the applicable 
requirements of Title 49 U.S.C. and this subchapter, an export 
airworthiness approval may be issued for a product or article located 
outside of the United States.



Sec. 21.327  Application.

    Any person may apply for an export airworthiness approval. Each 
applicant must apply in a form and manner prescribed by the FAA.



Sec. 21.329  Issuance of export certificates of airworthiness.

    (a) A person may obtain from the FAA an export certificate of 
airworthiness for an aircraft if--
    (1) A new or used aircraft manufactured under subpart F or G of this 
part meets the airworthiness requirements under subpart H of this part 
for a--
    (i) Standard airworthiness certificate; or
    (ii) Special airworthiness certificate in either the ``primary'' or 
the ``restricted'' category; or
    (2) A new or used aircraft not manufactured under subpart F or G of 
this part has a valid--
    (i) Standard airworthiness certificate; or
    (ii) Special airworthiness certificate in either the ``primary'' or 
the ``restricted'' category.
    (b) An aircraft need not meet a requirement specified in paragraph 
(a) of this section, as applicable, if--
    (1) The importing country or jurisdiction accepts, in a form and 
manner acceptable to the FAA, a deviation from that requirement; and
    (2) The export certificate of airworthiness lists as an exception 
any difference between the aircraft to be exported and its type design.



Sec. 21.331  Issuance of export airworthiness approvals for aircraft
engines, propellers, and articles.

    (a) A person may obtain from the FAA an export airworthiness 
approval to export a new aircraft engine, propeller, or article that is 
manufactured under this part if it conforms to its approved design and 
is in a condition for safe operation.
    (b) A new aircraft engine, propeller, or article need not meet a 
requirement of paragraph (a) of this section if--
    (1) The importing country or jurisdiction accepts, in a form and 
manner acceptable to the FAA, a deviation from that requirement; and

[[Page 167]]

    (2) The export airworthiness approval lists as an exception any 
difference between the aircraft engine, propeller, or article to be 
exported and its approved design.
    (c) A person may obtain from the FAA an export airworthiness 
approval to export a used aircraft engine, propeller, or article if it 
conforms to its approved design and is in a condition for safe 
operation.
    (d) A used aircraft engine or propeller need not meet a requirement 
of paragraph (c) of this section if--
    (1) The importing country or jurisdiction accepts, in a form and 
manner acceptable to the FAA, a deviation from that requirement; and
    (2) The export airworthiness approval lists as an exception any 
difference between the used aircraft engine or propeller to be exported 
and its approved design.



Sec. 21.335  Responsibilities of exporters.

    Unless otherwise agreed to by the importing country or jurisdiction, 
each exporter must--
    (a) Forward to the importing country or jurisdiction all documents 
specified by that country or jurisdiction;
    (b) Preserve and package products and articles as necessary to 
protect them against corrosion and damage during transit or storage and 
state the duration of effectiveness of such preservation and packaging;
    (c) Remove or cause to be removed any temporary installation 
incorporated on an aircraft for the purpose of export delivery and 
restore the aircraft to the approved configuration upon completion of 
the delivery flight;
    (d) Secure all proper foreign entry clearances from all the 
countries or jurisdictions involved when conducting sales demonstrations 
or delivery flights; and
    (e) When title to an aircraft passes or has passed to a foreign 
purchaser--
    (1) Request cancellation of the U.S. registration and airworthiness 
certificates from the FAA, giving the date of transfer of title, and the 
name and address of the foreign owner;
    (2) Return the Registration and Airworthiness Certificates to the 
FAA; and
    (3) Provide a statement to the FAA certifying that the U.S. 
identification and registration numbers have been removed from the 
aircraft in compliance with Sec. 45.33.

Subpart M [Reserved]



 Subpart N_Acceptance of Aircraft Engines, Propellers, and Articles for 
                                 Import

    Source: Docket No. FAA-2006-25877, 74 FR 53392, Amdt. 21-92, Oct. 
16, 2009, unless otherwise noted.



Sec. 21.500  Acceptance of aircraft engines and propellers.

    An aircraft engine or propeller manufactured in a foreign country or 
jurisdiction meets the requirements for acceptance under this subchapter 
if--
    (a) That country or jurisdiction is subject to the provisions of an 
agreement with the United States for the acceptance of that product;
    (b) That product is marked in accordance with part 45 of this 
chapter; and
    (c) The holder or licensee of a U.S. type certificate for that 
product furnishes with each such aircraft engine or propeller imported 
into the United States, an export airworthiness approval issued in 
accordance with the provisions of that agreement certifying that the 
individual aircraft engine or propeller--
    (1) Conforms to its U.S. type certificate and is in condition for 
safe operation; and
    (2) Has been subjected by the manufacturer to a final operational 
check.



Sec. 21.502  Acceptance of articles.

    An article (including an article produced under a letter of TSO 
design approval) manufactured in a foreign country or jurisdiction meets 
the requirements for acceptance under this subchapter if--
    (a) That country or jurisdiction is subject to the provisions of an 
agreement with the United States for the acceptance of that article;
    (b) That article is marked in accordance with part 45 of this 
chapter; and
    (c) An export airworthiness approval has been issued in accordance 
with the

[[Page 168]]

provisions of that agreement for that article for import into the United 
States.



              Subpart O_Technical Standard Order Approvals

    Source: Docket No. FAA-2006-25877, Amdt. 21-92, 74 FR 53392, Oct. 
16, 2009, unless otherwise noted.



Sec. 21.601  Applicability and definitions.

    (a) This subpart prescribes--
    (1) Procedural requirements for issuing TSO authorizations;
    (2) Rules governing the holders of TSO authorizations; and
    (3) Procedural requirements for issuing letters of TSO design 
approval.
    (b) For the purposes of this subpart--
    (1) A TSO issued by the FAA is a minimum performance standard for 
specified articles used on civil aircraft;
    (2) A TSO authorization is an FAA design and production approval 
issued to the manufacturer of an article that has been found to meet a 
specific TSO;
    (3) A letter of TSO design approval is an FAA design approval for an 
article that has been found to meet a specific TSO in accordance with 
the procedures of Sec. 21.621;
    (4) An article manufactured under a TSO authorization, an FAA letter 
of acceptance as described in Sec. 21.613(b), or an article 
manufactured under a letter of TSO design approval described in Sec. 
21.621 is an approved article for the purpose of meeting the regulations 
of this chapter that require the article to be approved; and
    (5) An article manufacturer is the person who controls the design 
and quality of the article produced (or to be produced, in the case of 
an application), including any related parts, processes, or services 
procured from an outside source.



Sec. 21.603  Application.

    (a) An applicant for a TSO authorization must apply to the 
appropriate aircraft certification office in the form and manner 
prescribed by the FAA. The applicant must include the following 
documents in the application:
    (1) A statement of conformance certifying that the applicant has met 
the requirements of this subpart and that the article concerned meets 
the applicable TSO that is effective on the date of application for that 
article.
    (2) One copy of the technical data required in the applicable TSO.
    (b) If the applicant anticipates a series of minor changes in 
accordance with Sec. 21.619, the applicant may set forth in its 
application the basic model number of the article and the part number of 
the components with open brackets after it to denote that suffix change 
letters or numbers (or combinations of them) will be added from time to 
time.
    (c) If the application is deficient, the applicant must, when 
requested by the FAA, provide any additional information necessary to 
show compliance with this part. If the applicant fails to provide the 
additional information within 30 days after the FAA's request, the FAA 
denies the application and notifies the applicant.



Sec. 21.605  Organization.

    Each applicant for or holder of a TSO authorization must provide the 
FAA with a document describing how the applicant's organization will 
ensure compliance with the provisions of this subpart. At a minimum, the 
document must describe assigned responsibilities and delegated 
authority, and the functional relationship of those responsible for 
quality to management and other organizational components.



Sec. 21.607  Quality system.

    Each applicant for or holder of a TSO authorization must establish a 
quality system that meets the requirements of Sec. 21.137.



Sec. 21.608  Quality manual.

    Each applicant for or holder of a TSO authorization must provide a 
manual describing its quality system to the FAA for approval. The manual 
must be in the English language and retrievable in a form acceptable to 
the FAA.



Sec. 21.609  Location of or change to manufacturing facilities.

    (a) An applicant may obtain a TSO authorization for manufacturing 
facilities located outside of the United

[[Page 169]]

States if the FAA finds no undue burden in administering the applicable 
requirements of Title 49 U.S.C. and this subchapter.
    (b) The TSO authorization holder must obtain FAA approval before 
making any changes to the location of any of its manufacturing 
facilities.
    (c) The TSO authorization holder must immediately notify the FAA, in 
writing, of any change to the manufacturing facilities that may affect 
the inspection, conformity, or airworthiness of its product or article.



Sec. 21.610  Inspections and tests.

    Each applicant for or holder of a TSO authorization must allow the 
FAA to inspect its quality system, facilities, technical data, and any 
manufactured articles and witness any tests, including any inspections 
or tests at a supplier facility, necessary to determine compliance with 
this subchapter.



Sec. 21.611  Issuance.

    If the FAA finds that the applicant complies with the requirements 
of this subchapter, the FAA issues a TSO authorization to the applicant 
(including all TSO deviations granted to the applicant).



Sec. 21.613  Duration.

    (a) A TSO authorization or letter of TSO design approval is 
effective until surrendered, withdrawn, or otherwise terminated by the 
FAA.
    (b) If a TSO is revised or canceled, the holder of an affected FAA 
letter of acceptance of a statement of conformance, TSO authorization, 
or letter of TSO design approval may continue to manufacture articles 
that meet the original TSO without obtaining a new acceptance, 
authorization, or approval but must comply with the requirements of this 
chapter.



Sec. 21.614  Transferability.

    The holder of a TSO authorization or letter of TSO design approval 
may not transfer the TSO authorization or letter of TSO design approval.



Sec. 21.616  Responsibility of holder.

    Each holder of a TSO authorization must--
    (a) Amend the document required by Sec. 21.605 as necessary to 
reflect changes in the organization and provide these amendments to the 
FAA.
    (b) Maintain a quality system in compliance with the data and 
procedures approved for the TSO authorization;
    (c) Ensure that each manufactured article conforms to its approved 
design, is in a condition for safe operation, and meets the applicable 
TSO;
    (d) Mark the TSO article for which an approval has been issued. 
Marking must be in accordance with part 45 of this chapter, including 
any critical parts;
    (e) Identify any portion of the TSO article (e.g., sub-assemblies, 
component parts, or replacement articles) that leave the manufacturer's 
facility as FAA approved with the manufacturer's part number and name, 
trademark, symbol, or other FAA approved manufacturer's identification;
    (f) Have access to design data necessary to determine conformity and 
airworthiness for each article produced under the TSO authorization. The 
manufacturer must retain this data until it no longer manufactures the 
article. At that time, copies of the data must be sent to the FAA;
    (g) Retain its TSO authorization and make it available to the FAA 
upon request; and
    (h) Make available to the FAA information regarding all delegation 
of authority to suppliers.



Sec. 21.618  Approval for deviation.

    (a) Each manufacturer who requests approval to deviate from any 
performance standard of a TSO must show that factors or design features 
providing an equivalent level of safety compensate for the standards 
from which a deviation is requested.
    (b) The manufacturer must send requests for approval to deviate, 
together with all pertinent data, to the appropriate aircraft 
certification office. If the article is manufactured under the authority 
of a foreign country or jurisdiction, the manufacturer must send 
requests for approval to deviate, together with all pertinent data, 
through the civil aviation authority of that country or jurisdiction to 
the FAA.

[[Page 170]]



Sec. 21.619  Design changes.

    (a) Minor changes by the manufacturer holding a TSO authorization. 
The manufacturer of an article under an authorization issued under this 
part may make minor design changes (any change other than a major 
change) without further approval by the FAA. In this case, the changed 
article keeps the original model number (part numbers may be used to 
identify minor changes) and the manufacturer must forward to the 
appropriate aircraft certification office, any revised data that are 
necessary for compliance with Sec. 21.603(b).
    (b) Major changes by the manufacturer holding a TSO authorization. 
Any design change by the manufacturer extensive enough to require a 
substantially complete investigation to determine compliance with a TSO 
is a major change. Before making a major change, the manufacturer must 
assign a new type or model designation to the article and apply for an 
authorization under Sec. 21.603.
    (c) Changes by persons other than the manufacturer. No design change 
by any person (other than the manufacturer who provided the statement of 
conformance for the article) is eligible for approval under this part 
unless the person seeking the approval is a manufacturer and applies 
under Sec. 21.603(a) for a separate TSO authorization. Persons other 
than a manufacturer may obtain approval for design changes under part 43 
or under the applicable airworthiness regulations of this chapter.



Sec. 21.620  Changes in quality system.

    After the issuance of a TSO authorization--
    (a) Each change to the quality system is subject to review by the 
FAA; and
    (b) The holder of the TSO authorization must immediately notify the 
FAA, in writing, of any change that may affect the inspection, 
conformity, or airworthiness of its article.



Sec. 21.621  Issue of letters of TSO design approval: Import articles.

    (a) The FAA may issue a letter of TSO design approval for an 
article--
    (1) Designed and manufactured in a foreign country or jurisdiction 
subject to the export provisions of an agreement with the United States 
for the acceptance of these articles for import; and
    (2) For import into the United States if--
    (i) The State of Design certifies that the article has been 
examined, tested, and found to meet the applicable TSO or the applicable 
performance standards of the State of Design and any other performance 
standards the FAA may prescribe to provide a level of safety equivalent 
to that provided by the TSO; and
    (ii) The manufacturer has provided to the FAA one copy of the 
technical data required in the applicable performance standard through 
its State of Design.
    (b) The FAA issues the letter of TSO design approval that lists any 
deviation granted under Sec. 21.618.

[Doc. No. FAA-2006-25877, Amdt. 21-92, 74 FR 53392, Oct. 16, 2009, as 
amended by Amdt. 21-92A, 75 FR 9095, Mar. 1, 2010]



             Subpart P_Special Federal Aviation Regulations

    Source: Docket No. FAA-2011-0186, Amdt. 21-92, 76 FR 12555, Mar. 8, 
2011, unless otherwise noted.



Sec. 21.700  SFAR No. 111--Lavatory Oxygen Systems.

    The requirements of Sec. 121.1500 of this chapter also apply to 
this part.



PART 23_AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC,
AND COMMUTER CATEGORY AIRPLANES--Table of Contents



Special Federal Aviation Regulation No. 23

                            Subpart A_General

Sec.
23.1 Applicability.
23.2 Special retroactive requirements.
23.3 Airplane categories.

                            Subpart B_Flight

                                 General

23.21 Proof of compliance.
23.23 Load distribution limits.
23.25 Weight limits.

[[Page 171]]

23.29 Empty weight and corresponding center of gravity.
23.31 Removable ballast.
23.33 Propeller speed and pitch limits.

                               Performance

23.45 General.
23.49 Stalling speed.
23.51 Takeoff speeds.
23.53 Takeoff performance.
23.55 Accelerate-stop distance.
23.57 Takeoff path.
23.59 Takeoff distance and takeoff run.
23.61 Takeoff flight path.
23.63 Climb: General.
23.65 Climb: All engines operating.
23.66 Takeoff climb: One-engine inoperative.
23.67 Climb: One engine inoperative.
23.69 Enroute climb/descent.
23.71 Glide: Single-engine airplanes.
23.73 Reference landing approach speed.
23.75 Landing distance.
23.77 Balked landing.

                         Flight Characteristics

23.141 General.

                   Controllability and Maneuverability

23.143 General.
23.145 Longitudinal control.
23.147 Directional and lateral control.
23.149 Minimum control speed.
23.151 Acrobatic maneuvers.
23.153 Control during landings.
23.155 Elevator control force in maneuvers.
23.157 Rate of roll.

                                  Trim

23.161 Trim.

                                Stability

23.171 General.
23.173 Static longitudinal stability.
23.175 Demonstration of static longitudinal stability.
23.177 Static directional and lateral stability.
23.181 Dynamic stability.

                                 Stalls

23.201 Wings level stall.
23.203 Turning flight and accelerated turning stalls.
23.207 Stall warning.

                                Spinning

23.221 Spinning.

                Ground and Water Handling Characteristics

23.231 Longitudinal stability and control.
23.233 Directional stability and control.
23.235 Operation on unpaved surfaces.
23.237 Operation on water.
23.239 Spray characteristics.

                    Miscellaneous Flight Requirements

23.251 Vibration and buffeting.
23.253 High speed characteristics.
23.255 Out of trim characteristics.

                           Subpart C_Structure

                                 General

23.301 Loads.
23.302 Canard or tandem wing configurations.
23.303 Factor of safety.
23.305 Strength and deformation.
23.307 Proof of structure.

                              Flight Loads

23.321 General.
23.331 Symmetrical flight conditions.
23.333 Flight envelope.
23.335 Design airspeeds.
23.337 Limit maneuvering load factors.
23.341 Gust loads factors.
23.343 Design fuel loads.
23.345 High lift devices.
23.347 Unsymmetrical flight conditions.
23.349 Rolling conditions.
23.351 Yawing conditions.
23.361 Engine torque.
23.363 Side load on engine mount.
23.365 Pressurized cabin loads.
23.367 Unsymmetrical loads due to engine failure.
23.369 Rear lift truss.
23.371 Gyroscopic and aerodynamic loads.
23.373 Speed control devices.

                    Control Surface and System Loads

23.391 Control surface loads.
23.393 Loads parallel to hinge line.
23.395 Control system loads.
23.397 Limit control forces and torques.
23.399 Dual control system.
23.405 Secondary control system.
23.407 Trim tab effects.
23.409 Tabs.
23.415 Ground gust conditions.

              Horizontal Stabilizing and Balancing Surfaces

23.421 Balancing loads.
23.423 Maneuvering loads.
23.425 Gust loads.
23.427 Unsymmetrical loads.

                            Vertical Surfaces

23.441 Maneuvering loads.
23.443 Gust loads.
23.445 Outboard fins or winglets.

                      Ailerons and Special Devices

23.455 Ailerons.
23.459 Special devices.

                              Ground Loads

23.471 General.

[[Page 172]]

23.473 Ground load conditions and assumptions.
23.477 Landing gear arrangement.
23.479 Level landing conditions.
23.481 Tail down landing conditions.
23.483 One-wheel landing conditions.
23.485 Side load conditions.
23.493 Braked roll conditions.
23.497 Supplementary conditions for tail wheels.
23.499 Supplementary conditions for nose wheels.
23.505 Supplementary conditions for ski-planes.
23.507 Jacking loads.
23.509 Towing loads.
23.511 Ground load; unsymmetrical loads on multiple-wheel units.

                               Water Loads

23.521 Water load conditions.
23.523 Design weights and center of gravity positions.
23.525 Application of loads.
23.527 Hull and main float load factors.
23.529 Hull and main float landing conditions.
23.531 Hull and main float takeoff condition.
23.533 Hull and main float bottom pressures.
23.535 Auxiliary float loads.
23.537 Seawing loads.

                      Emergency Landing Conditions

23.561 General.
23.562 Emergency landing dynamic conditions.

                           Fatigue Evaluation

23.571 Metallic pressurized cabin structures.
23.572 Metallic wing, empennage, and associated structures.
23.573 Damage tolerance and fatigue evaluation of structure.
23.574 Metallic damage tolerance and fatigue evaluation of commuter 
          category airplanes.
23.575 Inspections and other procedures.

                    Subpart D_Design and Construction

23.601 General.
23.603 Materials and workmanship.
23.605 Fabrication methods.
23.607 Fasteners.
23.609 Protection of structure.
23.611 Accessibility provisions.
23.613 Material strength properties and design values.
23.619 Special factors.
23.621 Casting factors.
23.623 Bearing factors.
23.625 Fitting factors.
23.627 Fatigue strength.
23.629 Flutter.

                                  Wings

23.641 Proof of strength.

                            Control Surfaces

23.651 Proof of strength.
23.655 Installation.
23.657 Hinges.
23.659 Mass balance.

                             Control Systems

23.671 General.
23.672 Stability augmentation and automatic and power-operated systems.
23.673 Primary flight controls.
23.675 Stops.
23.677 Trim systems.
23.679 Control system locks.
23.681 Limit load static tests.
23.683 Operation tests.
23.685 Control system details.
23.687 Spring devices.
23.689 Cable systems.
23.691 Artificial stall barrier system.
23.693 Joints.
23.697 Wing flap controls.
23.699 Wing flap position indicator.
23.701 Flap interconnection.
23.703 Takeoff warning system.

                              Landing Gear

23.721 General.
23.723 Shock absorption tests.
23.725 Limit drop tests.
23.726 Ground load dynamic tests.
23.727 Reserve energy absorption drop test.
23.729 Landing gear extension and retraction system.
23.731 Wheels.
23.733 Tires.
23.735 Brakes.
23.737 Skis.
23.745 Nose/tail wheel steering.

                            Floats and Hulls

23.751 Main float buoyancy.
23.753 Main float design.
23.755 Hulls.
23.757 Auxiliary floats.

                   Personnel and Cargo Accommodations

23.771 Pilot compartment.
23.773 Pilot compartment view.
23.775 Windshields and windows.
23.777 Cockpit controls.
23.779 Motion and effect of cockpit controls.
23.781 Cockpit control knob shape.
23.783 Doors.
23.785 Seats, berths, litters, safety belts, and shoulder harnesses.
23.787 Baggage and cargo compartments.
23.791 Passenger information signs.
23.803 Emergency evacuation.
23.805 Flightcrew emergency exits.
23.807 Emergency exits.
23.811 Emergency exit marking.

[[Page 173]]

23.812 Emergency lighting.
23.813 Emergency exit access.
23.815 Width of aisle.
23.831 Ventilation.

                             Pressurization

23.841 Pressurized cabins.
23.843 Pressurization tests.

                             Fire Protection

23.851 Fire extinguishers.
23.853 Passenger and crew compartment interiors.
23.855 Cargo and baggage compartment fire protection.
23.856 Thermal/acoustic insulation materials.
23.859 Combustion heater fire protection.
23.863 Flammable fluid fire protection.
23.865 Fire protection of flight controls, engine mounts, and other 
          flight structure.

               Electrical Bonding and Lightning Protection

23.867 Electrical bonding and protection against lightning and static 
          electricity.

                              Miscellaneous

23.871 Leveling means.

                          Subpart E_Powerplant

                                 General

23.901 Installation.
23.903 Engines.
23.904 Automatic power reserve system.
23.905 Propellers.
23.907 Propeller vibration and fatigue.
23.909 Turbocharger systems.
23.925 Propeller clearance.
23.929 Engine installation ice protection.
23.933 Reversing systems.
23.934 Turbojet and turbofan engine thrust reverser systems tests.
23.937 Turbopropeller-drag limiting systems.
23.939 Powerplant operating characteristics.
23.943 Negative acceleration.

                               Fuel System

23.951 General.
23.953 Fuel system independence.
23.954 Fuel system lightning protection.
23.955 Fuel flow.
23.957 Flow between interconnected tanks.
23.959 Unusable fuel supply.
23.961 Fuel system hot weather operation.
23.963 Fuel tanks: General.
23.965 Fuel tank tests.
23.967 Fuel tank installation.
23.969 Fuel tank expansion space.
23.971 Fuel tank sump.
23.973 Fuel tank filler connection.
23.975 Fuel tank vents and carburetor vapor vents.
23.977 Fuel tank outlet.
23.979 Pressure fueling systems.

                         Fuel System Components

23.991 Fuel pumps.
23.993 Fuel system lines and fittings.
23.994 Fuel system components.
23.995 Fuel valves and controls.
23.997 Fuel strainer or filter.
23.999 Fuel system drains.
23.1001 Fuel jettisoning system.

                               Oil System

23.1011 General.
23.1013 Oil tanks.
23.1015 Oil tank tests.
23.1017 Oil lines and fittings.
23.1019 Oil strainer or filter.
23.1021 Oil system drains.
23.1023 Oil radiators.
23.1027 Propeller feathering system.

                                 Cooling

23.1041 General.
23.1043 Cooling tests.
23.1045 Cooling test procedures for turbine engine powered airplanes.
23.1047 Cooling test procedures for reciprocating engine powered 
          airplanes.

                             Liquid Cooling

23.1061 Installation.
23.1063 Coolant tank tests.

                            Induction System

23.1091 Air induction system.
23.1093 Induction system icing protection.
23.1095 Carburetor deicing fluid flow rate.
23.1097 Carburetor deicing fluid system capacity.
23.1099 Carburetor deicing fluid system detail design.
23.1101 Induction air preheater design.
23.1103 Induction system ducts.
23.1105 Induction system screens.
23.1107 Induction system filters.
23.1109 Turbocharger bleed air system.
23.1111 Turbine engine bleed air system.

                             Exhaust System

23.1121 General.
23.1123 Exhaust system.
23.1125 Exhaust heat exchangers.

                   Powerplant Controls and Accessories

23.1141 Powerplant controls: General.
23.1142 Auxiliary power unit controls.
23.1143 Engine controls.
23.1145 Ignition switches.
23.1147 Mixture controls.
23.1149 Propeller speed and pitch controls.
23.1153 Propeller feathering controls.
23.1155 Turbine engine reverse thrust and propeller pitch settings below 
          the flight regime.

[[Page 174]]

23.1157 Carburetor air temperature controls.
23.1163 Powerplant accessories.
23.1165 Engine ignition systems.

                       Powerplant Fire Protection

23.1181 Designated fire zones; regions included.
23.1182 Nacelle areas behind firewalls.
23.1183 Lines, fittings, and components.
23.1189 Shutoff means.
23.1191 Firewalls.
23.1192 Engine accessory compartment diaphragm.
23.1193 Cowling and nacelle.
23.1195 Fire extinguishing systems.
23.1197 Fire extinguishing agents.
23.1199 Extinguishing agent containers.
23.1201 Fire extinguishing systems materials.
23.1203 Fire detector system.

                           Subpart F_Equipment

                                 General

23.1301 Function and installation.
23.1303 Flight and navigation instruments.
23.1305 Powerplant instruments.
23.1306 Electrical and electronic system lightning protection.
23.1307 Miscellaneous equipment.
23.1308 High-intensity Radiated Fields (HIRF) Protection.
23.1309 Equipment, systems, and installations.
23.1310 Power source capacity and distribution.

                        Instruments: Installation

23.1311 Electronic display instrument systems.
23.1321 Arrangement and visibility.
23.1322 Warning, caution, and advisory lights.
23.1323 Airspeed indicating system.
23.1325 Static pressure system.
23.1326 Pitot heat indication systems.
23.1327 Magnetic direction indicator.
23.1329 Automatic pilot system.
23.1331 Instruments using a power source.
23.1335 Flight director systems.
23.1337 Powerplant instruments installation.

                    Electrical Systems and Equipment

23.1351 General.
23.1353 Storage battery design and installation.
23.1357 Circuit protective devices.
23.1359 Electrical system fire protection.
23.1361 Master switch arrangement.
23.1365 Electric cables and equipment.
23.1367 Switches.

                                 Lights

23.1381 Instrument lights.
23.1383 Taxi and landing lights.
23.1385 Position light system installation.
23.1387 Position light system dihedral angles.
23.1389 Position light distribution and intensities.
23.1391 Minimum intensities in the horizontal plane of position lights.
23.1393 Minimum intensities in any vertical plane of position lights.
23.1395 Maximum intensities in overlapping beams of position lights.
23.1397 Color specifications.
23.1399 Riding light.
23.1401 Anticollision light system.

                            Safety Equipment

23.1411 General.
23.1415 Ditching equipment.
23.1416 Pneumatic de-icer boot system.
23.1419 Ice protection.

                         Miscellaneous Equipment

23.1431 Electronic equipment.
23.1435 Hydraulic systems.
23.1437 Accessories for multiengine airplanes.
23.1438 Pressurization and pneumatic systems.
23.1441 Oxygen equipment and supply.
23.1443 Minimum mass flow of supplemental oxygen.
23.1445 Oxygen distribution system.
23.1447 Equipment standards for oxygen dispensing units.
23.1449 Means for determining use of oxygen.
23.1450 Chemical oxygen generators.
23.1451 Fire protection for oxygen equipment.
23.1453 Protection of oxygen equipment from rupture.
23.1457 Cockpit voice recorders.
23.1459 Flight data recorders.
23.1461 Equipment containing high energy rotors.

             Subpart G_Operating Limitations and Information

23.1501 General.
23.1505 Airspeed limitations.
23.1507 Operating maneuvering speed.
23.1511 Flap extended speed.
23.1513 Minimum control speed.
23.1519 Weight and center of gravity.
23.1521 Powerplant limitations.
23.1522 Auxiliary power unit limitations.
23.1523 Minimum flight crew.
23.1524 Maximum passenger seating configuration.
23.1525 Kinds of operation.
23.1527 Maximum operating altitude.
23.1529 Instructions for Continued Airworthiness.

                          Markings and Placards

23.1541 General.

[[Page 175]]

23.1543 Instrument markings: General.
23.1545 Airspeed indicator.
23.1547 Magnetic direction indicator.
23.1549 Powerplant and auxiliary power unit instruments.
23.1551 Oil quantity indicator.
23.1553 Fuel quantity indicator.
23.1555 Control markings.
23.1557 Miscellaneous markings and placards.
23.1559 Operating limitations placard.
23.1561 Safety equipment.
23.1563 Airspeed placards.
23.1567 Flight maneuver placard.

           Airplane Flight Manual and Approved Manual Material

23.1581 General.
23.1583 Operating limitations.
23.1585 Operating procedures.
23.1587 Performance information.
23.1589 Loading information.

Appendix A to Part 23--Simplified Design Load Criteria
Appendix B to Part 23 [Reserved]
Appendix C to Part 23--Basic Landing Conditions
Appendix D to Part 23--Wheel Spin-Up and Spring-Back Loads
Appendix E to Part 23 [Reserved]
Appendix F to Part 23--Test Procedure
Appendix G to Part 23--Instructions for Continued Airworthiness
Appendix H to Part 23--Installation of An Automatic Power Reserve (APR) 
          System
Appendix I to Part 23--Seaplane Loads
Appendix J to Part 23--HIRF Environments and Equipment HIRF Test Levels

    Authority: 49 U.S.C. 106(g), 40113, 44701-44702, 44704.

    Source: Docket No. 4080, 29 FR 17955, Dec. 18. 1964; 30 FR 258, Jan. 
9, 1965, unless otherwise noted.



             Sec. Special Federal Aviation Regulation No. 23

    1. Applicability. An applicant is entitled to a type certificate in 
the normal category for a reciprocating or turbopropeller multiengine 
powered small airplane that is to be certificated to carry more than 10 
occupants and that is intended for use in operations under Part 135 of 
the Federal Aviation Regulations if he shows compliance with the 
applicable requirements of Part 23 of the Federal Aviation Regulations, 
as supplemented or modified by the additional airworthiness requirements 
of this regulation.
    2. References. Unless otherwise provided, all references in this 
regulation to specific sections of Part 23 of the Federal Aviation 
Regulations are those sections of Part 23 in effect on March 30, 1967.

                           Flight Requirements

    3. General. Compliance must be shown with the applicable 
requirements of Subpart B of Part 23 of the Federal Aviation Regulations 
in effect on March 30, 1967, as supplemented or modified in sections 4 
through 10 of this regulation.

                               Performance

    4. General. (a) Unless otherwise prescribed in this regulation, 
compliance with each applicable performance requirement in sections 4 
through 7 of this regulation must be shown for ambient atmospheric 
conditions and still air.
    (b) The performance must correspond to the propulsive thrust 
available under the particular ambient atmospheric conditions and the 
particular flight condition. The available propulsive thrust must 
correspond to engine power or thrust, not exceeding the approved power 
or thrust less--
    (1) Installation losses; and
    (2) The power or equivalent thrust absorbed by the accessories and 
services appropriate to the particular ambient atmospheric conditions 
and the particular flight condition.
    (c) Unless otherwise prescribed in this regulation, the applicant 
must select the take-off, en route, and landing configurations for the 
airplane.
    (d) The airplane configuration may vary with weight, altitude, and 
temperature, to the extent they are compatible with the operating 
procedures required by paragraph (e) of this section.
    (e) Unless otherwise prescribed in this regulation, in determining 
the critical engine inoperative takeoff performance, the accelerate-stop 
distance, takeoff distance, changes in the airplane's configuration, 
speed, power, and thrust, must be made in accordance with procedures 
established by the applicant for operation in service.
    (f) Procedures for the execution of balked landings must be 
established by the applicant and included in the Airplane Flight Manual.
    (g) The procedures established under paragraphs (e) and (f) of this 
section must--
    (1) Be able to be consistently executed in service by a crew of 
average skill;
    (2) Use methods or devices that are safe and reliable; and
    (3) Include allowance for any time delays, in the execution of the 
procedures, that may reasonably be expected in service.
    5. Takeoff--(a) General. The takeoff speeds described in paragraph 
(b), the accelerate-stop distance described in paragraph (c), and the 
takeoff distance described in paragraph (d), must be determined for--
    (1) Each weight, altitude, and ambient temperature within the 
operational limits selected by the applicant;

[[Page 176]]

    (2) The selected configuration for takeoff;
    (3) The center of gravity in the most unfavorable position;
    (4) The operating engine within approved operating limitation; and
    (5) Takeoff data based on smooth, dry, hard-surface runway.
    (b) Takeoff speeds. (1) The decision speed V1 is the 
calibrated airspeed on the ground at which, as a result of engine 
failure or other reasons, the pilot is assumed to have made a decision 
to continue or discontinue the takeoff. The speed V1 must be 
selected by the applicant but may not be less than--
    (i) 1.10 Vs1;
    (ii) 1.10 VMC;
    (iii) A speed that permits acceleration to V1 and stop in 
accordance with paragraph (c) allowing credit for an overrun distance 
equal to that required to stop the airplane from a ground speed of 35 
knots utilizing maximum braking; or
    (iv) A speed at which the airplane can be rotated for takeoff and 
shown to be adequate to safely continue the takeoff, using normal 
piloting skill, when the critical engine is suddenly made inoperative.
    (2) Other essential takeoff speeds necessary for safe operation of 
the airplane must be determined and shown in the Airplane Flight Manual.
    (c) Accelerate-stop distance. (1) The accelerate-stop distance is 
the sum of the distances necessary to--
    (i) Accelerate the airplane from a standing start to V1; 
and
    (ii) Decelerate the airplane from V1 to a speed not greater than 35 
knots, assuming that in the case of engine failure, failure of the 
critical engine is recognized by the pilot at the speed V1. 
The landing gear must remain in the extended position and maximum 
braking may be utilized during deceleration.
    (2) Means other than wheel brakes may be used to determine the 
accelerate-stop distance if that means is available with the critical 
engine inoperative and--
    (i) Is safe and reliable;
    (ii) Is used so that consistent results can be expected under normal 
operating conditions; and
    (iii) Is such that exceptional skill is not required to control the 
airplane.
    (d) All engines operating takeoff distance. The all engine operating 
takeoff distance is the horizontal distance required to takeoff and 
climb to a height of 50 feet above the takeoff surface according to 
procedures in FAR 23.51(a).
    (e) One-engine-inoperative takeoff. The maximum weight must be 
determined for each altitude and temperature within the operational 
limits established for the airplane, at which the airplane has takeoff 
capability after failure of the critical engine at or above 
V1 determined in accordance with paragraph (b) of this 
section. This capability may be established--
    (1) By demonstrating a measurably positive rate of climb with the 
airplane in the takeoff configuration, landing gear extended; or
    (2) By demonstrating the capability of maintaining flight after 
engine failure utilizing procedures prescribed by the applicant.
    6. Climb--(a) Landing climb: All-engines-operating. The maximum 
weight must be determined with the airplane in the landing 
configuration, for each altitude, and ambient temperature within the 
operational limits established for the airplane and with the most 
unfavorable center of gravity and out-of-ground effect in free air, at 
which the steady gradient of climb will not be less than 3.3 percent, 
with:
    (1) The engines at the power that is available 8 seconds after 
initiation of movement of the power or thrust controls from the mimimum 
flight idle to the takeoff position.
    (2) A climb speed not greater than the approach speed established 
under section 7 of this regulation and not less than the greater of 
1.05MC or 1.10VS1.
    (b) En route climb, one-engine-inoperative. (1) the maximum weight 
must be determined with the airplane in the en route configuration, the 
critical engine inoperative, the remaining engine at not more than 
maximum continuous power or thrust, and the most unfavorable center of 
gravity, at which the gradient at climb will be not less than--
    (i) 1.2 percent (or a gradient equivalent to 0.20 Vso2, if greater) 
at 5,000 feet and an ambient temperature of 41 [deg]F. or
    (ii) 0.6 percent (or a gradient equivalent to 0.01 Vso2, if greater) 
at 5,000 feet and ambient temperature of 81 [deg]F.
    (2) The minimum climb gradient specified in subdivisions (i) and 
(ii) of subparagraph (1) of this paragraph must vary linearly between 41 
[deg]F. and 81 [deg]F. and must change at the same rate up to the 
maximum operational temperature approved for the airplane.
    7. Landing. The landing distance must be determined for standard 
atmosphere at each weight and altitude in accordance with FAR 23.75(a), 
except that instead of the gliding approach specified in FAR 
23.75(a)(1), the landing may be preceded by a steady approach down to 
the 50-foot height at a gradient of descent not greater than 5.2 percent 
(3[deg]) at a calibrated airspeed not less than 1.3s1.

                                  Trim

    8. Trim--(a) Lateral and directional trim. The airplane must 
maintain lateral and directional trim in level flight at a speed of Vh 
or VMO/MMO, whichever is lower, with landing gear and wing flaps 
retracted.
    (b) Longitudinal trim. The airplane must maintain longitudinal trim 
during the following conditions, except that it need not

[[Page 177]]

maintain trim at a speed greater than VMO/MMO:
    (1) In the approach conditions specified in FAR 23.161(c)(3) through 
(5), except that instead of the speeds specified therein, trim must be 
maintained with a stick force of not more than 10 pounds down to a speed 
used in showing compliance with section 7 of this regulation or 1.4 Vs1 
whichever is lower.
    (2) In level flight at any speed from VH or VMO/MMO, whichever is 
lower, to either Vx or 1.4 Vs1, with the landing gear and 
wing flaps retracted.

                                Stability

    9. Static longitudinal stability. (a) In showing compliance with the 
provisions of FAR 23.175(b) and with paragraph (b) of this section, the 
airspeed must return to within 7\1/2\ percent of 
the trim speed.
    (b) Cruise stability. The stick force curve must have a stable slope 
for a speed range of 50 knots from the trim speed 
except that the speeds need not exceed VFC/MFC or be less than 1.4 
Vs1. This speed range will be considered to begin at the 
outer extremes of the friction band and the stick force may not exceed 
50 pounds with--
    (i) Landing gear retracted;
    (ii) Wing flaps retracted;
    (iii) The maximum cruising power as selected by the applicant as an 
operating limitation for turbine engines or 75 percent of maximum 
continuous power for reciprocating engines except that the power need 
not exceed that required at VMO/MMO:
    (iv) Maximum takeoff weight; and
    (v) The airplane trimmed for level flight with the power specified 
in subparagraph (iii) of this paragraph.
    VFC/MFC may not be less than a speed midway between VMO/MMO and VDF/
MDF, except that, for altitudes where Mach number is the limiting 
factor, MFC need not exceed the Mach number at which effective speed 
warning occurs.
    (c) Climb stability. For turbopropeller powered airplanes only. In 
showing compliance with FAR 23.175(a), an applicant must in lieu of the 
power specified in FAR 23.175(a)(4), use the maximum power or thrust 
selected by the applicant as an operating limitation for use during 
climb at the best rate of climb speed except that the speed need not be 
less than 1.4 Vs1.

                                 Stalls

    10. Stall warning. If artificial stall warning is required to comply 
with the requirements of FAR 23.207, the warning device must give 
clearly distinguishable indications under expected conditions of flight. 
The use of a visual warning device that requires the attention of the 
crew within the cockpit is not acceptable by itself.

                             Control Systems

    11. Electric trim tabs. The airplane must meet the requirements of 
FAR 23.677 and in addition it must be shown that the airplane is safely 
controllable and that a pilot can perform all the maneuvers and 
operations necessary to effect a safe landing following any probable 
electric trim tab runaway which might be reasonably expected in service 
allowing for appropriate time delay after pilot recognition of the 
runaway. This demonstration must be conducted at the critical airplane 
weights and center of gravity positions.

                        Instruments: Installation

    12. Arrangement and visibility. Each instrument must meet the 
requirements of FAR 23.1321 and in addition--
    (a) Each flight, navigation, and powerplant instrument for use by 
any pilot must be plainly visible to him from his station with the 
minimum practicable deviation from his normal position and line of 
vision when he is looking forward along the flight path.
    (b) The flight instruments required by FAR 23.1303 and by the 
applicable operating rules must be grouped on the instrument panel and 
centered as nearly as practicable about the vertical plane of each 
pilot's forward vision. In addition--
    (1) The instrument that most effectively indicates the attitude must 
be on the panel in the top center position;
    (2) The instrument that most effectively indicates airspeed must be 
adjacent to and directly to the left of the instrument in the top center 
position;
    (3) The instrument that most effectively indicates altitude must be 
adjacent to and directly to the right of the instrument in the top 
center position; and
    (4) The instrument that most effectively indicates direction of 
flight must be adjacent to and directly below the instrument in the top 
center position.
    13. Airspeed indicating system. Each airspeed indicating system must 
meet the requirements of FAR 23.1323 and in addition--
    (a) Airspeed indicating instruments must be of an approved type and 
must be calibrated to indicate true airspeed at sea level in the 
standard atmosphere with a mimimum practicable instrument calibration 
error when the corresponding pilot and static pressures are supplied to 
the instruments.
    (b) The airspeed indicating system must be calibrated to determine 
the system error, i.e., the relation between IAS and CAS, in flight and 
during the accelerate takeoff ground run. The ground run calibration 
must be obtained between 0.8 of the mimimum value of V1 and 
1.2 times the maximum value

[[Page 178]]

of V1, considering the approved ranges of altitude and 
weight. The ground run calibration will be determined assuming an engine 
failure at the mimimum value of V1.
    (c) The airspeed error of the installation excluding the instrument 
calibration error, must not exceed 3 percent or 5 knots whichever is 
greater, throughout the speed range from VMO to 1.3S1 with 
flaps retracted and from 1.3 VSO to VFE with flaps in the 
landing position.
    (d) Information showing the relationship between IAS and CAS must be 
shown in the Airplane Flight Manual.
    14. Static air vent system. The static air vent system must meet the 
requirements of FAR 23.1325. The altimeter system calibration must be 
determined and shown in the Airplane Flight Manual.

                  Operating Limitations and Information

    15. Maximum operating limit speed VMO/MMO. Instead of establishing 
operating limitations based on VME and VNO, the applicant must establish 
a maximum operating limit speed VMO/MMO in accordance with the 
following:
    (a) The maximum operating limit speed must not exceed the design 
cruising speed Vc and must be sufficiently below VD/MD or VDF/MDF to 
make it highly improbable that the latter speeds will be inadvertently 
exceeded in flight.
    (b) The speed Vmo must not exceed 0.8 VD/MD or 0.8 VDF/MDF unless 
flight demonstrations involving upsets as specified by the Administrator 
indicates a lower speed margin will not result in speeds exceeding VD/MD 
or VDF. Atmospheric variations, horizontal gusts, and equipment errors, 
and airframe production variations will be taken into account.
    16. Minimum flight crew. In addition to meeting the requirements of 
FAR 23.1523, the applicant must establish the minimum number and type of 
qualified flight crew personnel sufficient for safe operation of the 
airplane considering--
    (a) Each kind of operation for which the applicant desires approval;
    (b) The workload on each crewmember considering the following:
    (1) Flight path control.
    (2) Collision avoidance.
    (3) Navigation.
    (4) Communications.
    (5) Operation and monitoring of all essential aircraft systems.
    (6) Command decisions; and
    (c) The accessibility and ease of operation of necessary controls by 
the appropriate crewmember during all normal and emergency operations 
when at his flight station.
    17. Airspeed indicator. The airspeed indicator must meet the 
requirements of FAR 23.1545 except that, the airspeed notations and 
markings in terms of VNO and VNE must be replaced by the VMO/MMO 
notations. The airspeed indicator markings must be easily read and 
understood by the pilot. A placard adjacent to the airspeed indicator is 
an acceptable means of showing compliance with the requirements of FAR 
23.1545(c).

                         Airplane Flight Manual

    18. General. The Airplane Flight Manual must be prepared in 
accordance with the requirements of FARs 23.1583 and 23.1587, and in 
addition the operating limitations and performance information set forth 
in sections 19 and 20 must be included.
    19. Operating limitations. The Airplane Flight Manual must include 
the following limitations--
    (a) Airspeed limitations. (1) The maximum operating limit speed VMO/
MMO and a statement that this speed limit may not be deliberately 
exceeded in any regime of flight (climb, cruise, or descent) unless a 
higher speed is authorized for flight test or pilot training;
    (2) If an airspeed limitation is based upon compressibility effects, 
a statement to this effect and information as to any symptoms, the 
probable behavior of the airplane, and the recommended recovery 
procedures; and
    (3) The airspeed limits, shown in terms of VMO/MMO instead of VNO 
and VNE.
    (b) Takeoff weight limitations. The maximum takeoff weight for each 
airport elevation, ambient temperature, and available takeoff runway 
length within the range selected by the applicant. This weight may not 
exceed the weight at which:
    (1) The all-engine operating takeoff distance determined in 
accordance with section 5(d) or the accelerate-stop distance determined 
in accordance with section 5(c), which ever is greater, is equal to the 
available runway length;
    (2) The airplane complies with the one-engine-inoperative takeoff 
requirements specified in section 5(e); and
    (3) The airplane complies with the one-engine-inoperative en route 
climb requirements specified in section 6(b), assuming that a standard 
temperature lapse rate exists from the airport elevation to the altitude 
of 5,000 feet, except that the weight may not exceed that corresponding 
to a temperature of 41 [deg]F at 5,000 feet.
    20. Performance information. The Airplane Flight Manual must contain 
the performance information determined in accordance with the provisions 
of the performance requirements of this regulation. The information must 
include the following:
    (a) Sufficient information so that the take-off weight limits 
specified in section 19(b) can be determined for all temperatures and 
altitudes within the operation limitations selected by the applicant.

[[Page 179]]

    (b) The conditions under which the performance information was 
obtained, including the airspeed at the 50-foot height used to determine 
landing distances.
    (c) The performance information (determined by extrapolation and 
computed for the range of weights between the maximum landing and 
takeoff weights) for--
    (1) Climb in the landing configuration; and
    (2) Landing distance.
    (d) Procedure established under section 4 of this regulation related 
to the limitations and information required by this section in the form 
of guidance material including any relevant limitations or information.
    (e) An explanation of significant or unusual flight or ground 
handling characteristics of the airplane.
    (f) Airspeeds, as indicated airspeeds, corresponding to those 
determined for takeoff in accordance with section 5(b).
    21. Maximum operating altitudes. The maximum operating altitude to 
which operation is permitted, as limited by flight, structural, 
powerplant, functional, or equipment characteristics, must be specified 
in the Airplane Flight Manual.
    22. Stowage provision for Airplane Flight Manual. Provision must be 
made for stowing the Airplane Flight Manual in a suitable fixed 
container which is readily accessible to the pilot.
    23. Operating procedures. Procedures for restarting turbine engines 
in flight (including the effects of altitude) must be set forth in the 
Airplane Flight Manual.

                          Airframe Requirements

                              flight loads

    24. Engine torque. (a) Each turbopropeller engine mount and its 
supporting structure must be designed for the torque effects of--
    (1) The conditions set forth in FAR 23.361(a).
    (2) The limit engine torque corresponding to takeoff power and 
propeller speed, multiplied by a factor accounting for propeller control 
system malfunction, including quick feathering action, simultaneously 
with 1 g level flight loads. In the absence of a rational analysis, a 
factor of 1.6 must be used.
    (b) The limit torque is obtained by multiplying the mean torque by a 
factor of 1.25.
    25. Turbine engine gyroscopic loads. Each turbopropeller engine 
mount and its supporting structure must be designed for the gyroscopic 
loads that result, with the engines at maximum continuous r.p.m., under 
either--
    (a) The conditions prescribed in FARs 23.351 and 23.423; or
    (b) All possible combinations of the following:
    (1) A yaw velocity of 2.5 radius per second.
    (2) A pitch velocity of 1.0 radians per second.
    (3) A normal load factor of 2.5.
    (4) Maximum continuous thrust.
    26. Unsymmetrical loads due to engine failure. (a) Turbopropeller 
powered airplanes must be designed for the unsymmetrical loads resulting 
from the failure of the critical engine including the following 
conditions in combination with a single malfunction of the propeller 
drag limiting system, considering the probable pilot corrective action 
on the flight controls.
    (1) At speeds between VMC and VD, the loads resulting from power 
failure because of fuel flow interruption are considered to be limit 
loads.
    (2) At speeds between VMC and VC, the loads resulting from the 
disconnection of the engine compressor from the turbine or from loss of 
the turbine blades are considered to be ultimate loads.
    (3) The time history of the thrust decay and drag buildup occurring 
as a result of the prescribed engine failures must be substantiated by 
test or other data applicable to the particular engine-propeller 
combination.
    (4) The timing and magnitude of the probable pilot corrective action 
must be conservatively estimated, considering the characteristics of the 
particular engine-propeller-airplane combination.
    (b) Pilot corrective action may be assumed to be initiated at the 
time maximum yawing velocity is reached, but not earlier than two 
seconds after the engine failure. The magnitude of the corrective action 
may be based on the control forces specified in FAR 23.397 except that 
lower forces may be assumed where it is shown by analysis or test that 
these forces can control the yaw and roll resulting from the prescribed 
engine failure conditions.

                              Ground Loads

    27. Dual wheel landing gear units. Each dual wheel landing gear unit 
and its supporting structure must be shown to comply with the following:
    (a) Pivoting. The airplane must be assumed to pivot about one side 
of the main gear with the brakes on that side locked. The limit vertical 
load factor must be 1.0 and the coefficient of friction 0.8. This 
condition need apply only to the main gear and its supporting structure.
    (b) Unequal tire inflation. A 60-40 percent distribution of the 
loads established in accordance with FAR 23.471 through FAR 23.483 must 
be applied to the dual wheels.
    (c) Flat tire. (1) Sixty percent of the loads specified in FAR 
23.471 through FAR 23.483 must be applied to either wheel in a unit.
    (2) Sixty percent of the limit drag and side loads and 100 percent 
of the limit vertical load established in accordance with FARs 23.493 
and 23.485 must be applied to either wheel in a unit except that the 
vertical load

[[Page 180]]

need not exceed the maximum vertical load in paragraph (c)(1) of this 
section.

                           Fatigue Evaluation

    28. Fatigue evaluation of wing and associated structure. Unless it 
is shown that the structure, operating stress levels, materials, and 
expected use are comparable from a fatigue standpoint to a similar 
design which has had substantial satisfactory service experience, the 
strength, detail design, and the fabrication of those parts of the wing, 
wing carrythrough, and attaching structure whose failure would be 
catastrophic must be evaluated under either--
    (a) A fatigue strength investigation in which the structure is shown 
by analysis, tests, or both to be able to withstand the repeated loads 
of variable magnitude expected in service; or
    (b) A fail-safe strength investigation in which it is shown by 
analysis, tests, or both that catastrophic failure of the structure is 
not probable after fatigue, or obvious partial failure, of a principal 
structural element, and that the remaining structure is able to 
withstand a static ultimate load factor of 75 percent of the critical 
limit load factor at Vc. These loads must be multiplied by a factor of 
1.15 unless the dynamic effects of failure under static load are 
otherwise considered.

                         Design and Construction

    29. Flutter. For Multiengine turbopropeller powered airplanes, a 
dynamic evaluation must be made and must include--
    (a) The significant elastic, inertia, and aerodynamic forces 
associated with the rotations and displacements of the plane of the 
propeller; and
    (b) Engine-propeller-nacelle stiffness and damping variations 
appropriate to the particular configuration.

                              Landing Gear

    30. Flap operated landing gear warning device. Airplanes having 
retractable landing gear and wing flaps must be equipped with a warning 
device that functions continuously when the wing flaps are extended to a 
flap position that activates the warning device to give adequate warning 
before landing, using normal landing procedures, if the landing gear is 
not fully extended and locked. There may not be a manual shut off for 
this warning device. The flap position sensing unit may be installed at 
any suitable location. The system for this device may use any part of 
the system (including the aural warning device) provided for other 
landing gear warning devices.

                   Personnel and Cargo Accommodations

    31. Cargo and baggage compartments. Cargo and baggage compartments 
must be designed to meet the requirements of FAR 23.787 (a) and (b), and 
in addition means must be provided to protect passengers from injury by 
the contents of any cargo or baggage compartment when the ultimate 
forward inertia force is 9g.
    32. Doors and exits. The airplane must meet the requirements of FAR 
23.783 and FAR 23.807 (a)(3), (b), and (c), and in addition:
    (a) There must be a means to lock and safeguard each external door 
and exit against opening in flight either inadvertently by persons, or 
as a result of mechanical failure. Each external door must be operable 
from both the inside and the outside.
    (b) There must be means for direct visual inspection of the locking 
mechanism by crewmembers to determine whether external doors and exits, 
for which the initial opening movement is outward, are fully locked. In 
addition, there must be a visual means to signal to crewmembers when 
normally used external doors are closed and fully locked.
    (c) The passenger entrance door must qualify as a floor level 
emergency exit. Each additional required emergency exit except floor 
level exits must be located over the wing or must be provided with 
acceptable means to assist the occupants in descending to the ground. In 
addition to the passenger entrance door:
    (1) For a total seating capacity of 15 or less, an emergency exit as 
defined in FAR 23.807(b) is required on each side of the cabin.
    (2) For a total seating capacity of 16 through 23, three emergency 
exits as defined in 23.807(b) are required with one on the same side as 
the door and two on the side opposite the door.
    (d) An evacuation demonstration must be conducted utilizing the 
maximum number of occupants for which certification is desired. It must 
be conducted under simulated night conditions utilizing only the 
emergency exits on the most critical side of the aircraft. The 
participants must be representative of average airline passengers with 
no prior practice or rehearsal for the demonstration. Evacuation must be 
completed within 90 seconds.
    (e) Each emergency exit must be marked with the word ``Exit'' by a 
sign which has white letters 1 inch high on a red background 2 inches 
high, be self-illuminated or independently internally electrically 
illuminated, and have a minimum luminescence (brightness) of at least 
160 microlamberts. The colors may be reversed if the passenger 
compartment illumination is essentially the same.
    (f) Access to window type emergency exits must not be obstructed by 
seats or seat backs.
    (g) The width of the main passenger aisle at any point between seats 
must equal or exceed the values in the following table.

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------------------------------------------------------------------------
                                    Minimum main passenger aisle width
                                 ---------------------------------------
     Total seating capacity          Less than 25     25 inches and more
                                   inches from floor      from floor
------------------------------------------------------------------------
10 through 23...................  9 inches..........  15 inches.
------------------------------------------------------------------------

                              Miscellaneous

    33. Lightning strike protection. Parts that are electrically 
insulated from the basic airframe must be connected to it through 
lightning arrestors unless a lightning strike on the insulated part--
    (a) Is improbable because of shielding by other parts; or
    (b) Is not hazardous.
    34. Ice protection. If certification with ice protection provisions 
is desired, compliance with the following requirements must be shown:
    (a) The recommended procedures for the use of the ice protection 
equipment must be set forth in the Airplane Flight Manual.
    (b) An analysis must be performed to establish, on the basis of the 
airplane's operational needs, the adequacy of the ice protection system 
for the various components of the airplane. In addition, tests of the 
ice protection system must be conducted to demonstrate that the airplane 
is capable of operating safely in continuous maximum and intermittent 
maximum icing conditions as described in FAR 25, appendix C.
    (c) Compliance with all or portions of this section may be 
accomplished by reference, where applicable because of similarity of the 
designs, to analysis and tests performed by the applicant for a type 
certificated model.
    35. Maintenance information. The applicant must make available to 
the owner at the time of delivery of the airplane the information he 
considers essential for the proper maintenance of the airplane. That 
information must include the following:
    (a) Description of systems, including electrical, hydraulic, and 
fuel controls.
    (b) Lubrication instructions setting forth the frequency and the 
lubricants and fluids which are to be used in the various systems.
    (c) Pressures and electrical loads applicable to the various 
systems.
    (d) Tolerances and adjustments necessary for proper functioning.
    (e) Methods of leveling, raising, and towing.
    (f) Methods of balancing control surfaces.
    (g) Identification of primary and secondary structures.
    (h) Frequency and extent of inspections necessary to the proper 
operation of the airplane.
    (i) Special repair methods applicable to the airplane.
    (j) Special inspection techniques, including those that require X-
ray, ultrasonic, and magnetic particle inspection.
    (k) List of special tools.

                               Propulsion

                                 general

    36. Vibration characteristics. For turbopropeller powered airplanes, 
the engine installation must not result in vibration characteristics of 
the engine exceeding those established during the type certification of 
the engine.
    37. In-flight restarting of engine. If the engine on turbopropeller 
powered airplanes cannot be restarted at the maximum cruise altitude, a 
determination must be made of the altitude below which restarts can be 
consistently accomplished. Restart information must be provided in the 
Airplane Flight Manual.
    38. Engines--(a) For turbopropeller powered airplanes. The engine 
installation must comply with the following requirements:
    (1) Engine isolation. The powerplants must be arranged and isolated 
from each other to allow operation, in at least one configuration, so 
that the failure or malfunction of any engine, or of any system that can 
affect the engine, will not--
    (i) Prevent the continued safe operation of the remaining engines; 
or
    (ii) Require immediate action by any crewmember for continued safe 
operation.
    (2) Control of engine rotation. There must be a means to 
individually stop and restart the rotation of any engine in flight 
except that engine rotation need not be stopped if continued rotation 
could not jeopardize the safety of the airplane. Each component of the 
stopping and restarting system on the engine side of the firewall, and 
that might be exposed to fire, must be at least fire resistant. If 
hydraulic propeller feathering systems are used for this purpose, the 
feathering lines must be at least fire resistant under the operating 
conditions that may be expected to exist during feathering.
    (3) Engine speed and gas temperature control devices. The powerplant 
systems associated with engine control devices, systems, and 
instrumentation must provide reasonable assurance that those engine 
operating limitations that adversely affect turbine rotor structural 
integrity will not be exceeded in service.
    (b) For reciprocating-engine powered airplanes. To provide engine 
isolation, the powerplants must be arranged and isolated from each other 
to allow operation, in at least one configuration, so that the failure 
or malfunction of any engine, or of any system that can affect that 
engine, will not--
    (1) Prevent the continued safe operation of the remaining engines; 
or
    (2) Require immediate action by any crewmember for continued safe 
operation.
    39. Turbopropeller reversing systems. (a) Turbopropeller reversing 
systems intended for

[[Page 182]]

ground operation must be designed so that no single failure or 
malfunction of the system will result in unwanted reverse thrust under 
any expected operating condition. Failure of structural elements need 
not be considered if the probability of this kind of failure is 
extremely remote.
    (b) Turbopropeller reversing systems intended for in-flight use must 
be designed so that no unsafe condition will result during normal 
operation of the system, or from any failure (or reasonably likely 
combination of failures) of the reversing system, under any anticipated 
condition of operation of the airplane. Failure of structural elements 
need not be considered if the probability of this kind of failure is 
extremely remote.
    (c) Compliance with this section may be shown by failure analysis, 
testing, or both for propeller systems that allow propeller blades to 
move from the flight low-pitch position to a position that is 
substantially less than that at the normal flight low-pitch stop 
position. The analysis may include or be supported by the analysis made 
to show compliance with the type certification of the propeller and 
associated installation components. Credit will be given for pertinent 
analysis and testing completed by the engine and propeller 
manufacturers.
    40. Turbopropeller drag-limiting systems. Turbopropeller drag-
limiting systems must be designed so that no single failure or 
malfunction of any of the systems during normal or emergency operation 
results in propeller drag in excess of that for which the airplane was 
designed. Failure of structural elements of the drag-limiting systems 
need not be considered if the probability of this kind of failure is 
extremely remote.
    41. Turbine engine powerplant operating characteristics. For 
turbopropeller powered airplanes, the turbine engine powerplant 
operating characteristics must be investigated in flight to determine 
that no adverse characteristics (such as stall, surge, or flameout) are 
present to a hazardous degree, during normal and emergency operation 
within the range of operating limitations of the airplane and of the 
engine.
    42. Fuel flow. (a) For turbopropeller powered airplanes--
    (1) The fuel system must provide for continuous supply of fuel to 
the engines for normal operation without interruption due to depletion 
of fuel in any tank other than the main tank; and
    (2) The fuel flow rate for turbopropeller engine fuel pump systems 
must not be less than 125 percent of the fuel flow required to develop 
the standard sea level atmospheric conditions takeoff power selected and 
included as an operating limitation in the Airplane Flight Manual.
    (b) For reciprocating engine powered airplanes, it is acceptable for 
the fuel flow rate for each pump system (main and reserve supply) to be 
125 percent of the takeoff fuel consumption of the engine.

                         Fuel System Components

    43. Fuel pumps. For turbopropeller powered airplanes, a reliable and 
independent power source must be provided for each pump used with 
turbine engines which do not have provisions for mechanically driving 
the main pumps. It must be demonstrated that the pump installations 
provide a reliability and durability equivalent to that provided by FAR 
23.991(a).
    44. Fuel strainer or filter. For turbopropeller powered airplanes, 
the following apply:
    (a) There must be a fuel strainer or filter between the tank outlet 
and the fuel metering device of the engine. In addition, the fuel 
strainer or filter must be--
    (1) Between the tank outlet and the engine-driven positive 
displacement pump inlet, if there is an engine-driven positive 
displacement pump;
    (2) Accessible for drainage and cleaning and, for the strainer 
screen, easily removable; and
    (3) Mounted so that its weight is not supported by the connecting 
lines or by the inlet or outlet connections of the strainer or filter 
itself.
    (b) Unless there are means in the fuel system to prevent the 
accumulation of ice on the filter, there must be means to automatically 
maintain the fuel flow if ice-clogging of the filter occurs; and
    (c) The fuel strainer or filter must be of adequate capacity (with 
respect to operating limitations established to insure proper service) 
and of appropriate mesh to insure proper engine operation, with the fuel 
contaminated to a degree (with respect to particle size and density) 
that can be reasonably expected in service. The degree of fuel filtering 
may not be less than that established for the engine type certification.
    45. Lightning strike protection. Protection must be provided against 
the ignition of flammable vapors in the fuel vent system due to 
lightning strikes.

                                 Cooling

    46. Cooling test procedures for turbopropeller powered airplanes. 
(a) Turbopropeller powered airplanes must be shown to comply with the 
requirements of FAR 23.1041 during takeoff, climb en route, and landing 
stages of flight that correspond to the applicable performance 
requirements. The cooling test must be conducted with the airplane in 
the configuration and operating under the conditions that are critical 
relative to cooling during each stage of flight. For the cooling tests a 
temperature is ``stabilized'' when its rate of change is less than 2 
[deg]F. per minute.
    (b) Temperatures must be stabilized under the conditions from which 
entry is made into

[[Page 183]]

each stage of flight being investigated unless the entry condition is 
not one during which component and engine fluid temperatures would 
stabilize, in which case, operation through the full entry condition 
must be conducted before entry into the stage of flight being 
investigated in order to allow temperatures to reach their natural 
levels at the time of entry. The takeoff cooling test must be preceded 
by a period during which the powerplant component and engine fluid 
temperatures are stabilized with the engines at ground idle.
    (c) Cooling tests for each stage of flight must be continued until--
    (1) The component and engine fluid temperatures stabilize;
    (2) The stage of flight is completed; or
    (3) An operating limitation is reached.

                            Induction System

    47. Air induction. For turbopropeller powered airplanes--
    (a) There must be means to prevent hazardous quantities of fuel 
leakage or overflow from drains, vents, or other components of flammable 
fluid systems from entering the engine intake system; and
    (b) The air inlet ducts must be located or protected so as to 
minimize the ingestion of foreign matter during takeoff, landing, and 
taxiing.
    48. Induction system icing protection. For turbopropeller powered 
airplanes, each turbine engine must be able to operate throughout its 
flight power range without adverse effect on engine operation or serious 
loss of power or thrust, under the icing conditions specified in 
appendix C of FAR 25. In addition, there must be means to indicate to 
appropriate flight crewmembers the functioning of the powerplant ice 
protection system.
    49. Turbine engine bleed air systems. Turbine engine bleed air 
systems of turbopropeller powered airplanes must be investigated to 
determine--
    (a) That no hazard to the airplane will result if a duct rupture 
occurs. This condition must consider that a failure of the duct can 
occur anywhere between the engine port and the airplane bleed service; 
and
    (b) That if the bleed air system is used for direct cabin 
pressurization, it is not possible for hazardous contamination of the 
cabin air system to occur in event of lubrication system failure.

                             Exhaust System

    50. Exhaust system drains. Turbopropeller engine exhaust systems 
having low spots or pockets must incorporate drains at such locations. 
These drains must discharge clear of the airplane in normal and ground 
attitudes to prevent the accumulation of fuel after the failure of an 
attempted engine start.

                   Powerplant Controls and Accessories

    51. Engine controls. If throttles or power levers for turbopropeller 
powered airplanes are such that any position of these controls will 
reduce the fuel flow to the engine(s) below that necessary for 
satisfactory and safe idle operation of the engine while the airplane is 
in flight, a means must be provided to prevent inadvertent movement of 
the control into this position. The means provided must incorporate a 
positive lock or stop at this idle position and must require a separate 
and distinct operation by the crew to displace the control from the 
normal engine operating range.
    52. Reverse thrust controls. For turbopropeller powered airplanes, 
the propeller reverse thrust controls must have a means to prevent their 
inadvertent operation. The means must have a positive lock or stop at 
the idle position and must require a separate and distinct operation by 
the crew to displace the control from the flight regime.
    53. Engine ignition systems. Each turbopropeller airplane ignition 
system must be considered an essential electrical load.
    54. Powerplant accessories. The powerplant accessories must meet the 
requirements of FAR 23.1163, and if the continued rotation of any 
accessory remotely driven by the engine is hazardous when malfunctioning 
occurs, there must be means to prevent rotation without interfering with 
the continued operation of the engine.

                       Powerplant Fire Protection

    55. Fire detector system. For turbopropeller powered airplanes, the 
following apply:
    (a) There must be a means that ensures prompt detection of fire in 
the engine compartment. An overtemperature switch in each engine cooling 
air exit is an acceptable method of meeting this requirement.
    (b) Each fire detector must be constructed and installed to 
withstand the vibration, inertia, and other loads to which it may be 
subjected in operation.
    (c) No fire detector may be affected by any oil, water, other 
fluids, or fumes that might be present.
    (d) There must be means to allow the flight crew to check, in 
flight, the functioning of each fire detector electric circuit.
    (e) Wiring and other components of each fire detector system in a 
fire zone must be at least fire resistant.
    56. Fire protection, cowling and nacelle skin. For reciprocating 
engine powered airplanes, the engine cowling must be designed and 
constructed so that no fire originating in the engine compartment can 
enter, either through openings or by burn through, any other region 
where it would create additional hazards.
    57. Flammable fluid fire protection. If flammable fluids or vapors 
might be liberated by

[[Page 184]]

the leakage of fluid systems in areas other than engine compartments, 
there must be means to--
    (a) Prevent the ignition of those fluids or vapors by any other 
equipment; or
    (b) Control any fire resulting from that ignition.

                                Equipment

    58. Powerplant instruments. (a) The following are required for 
turbopropeller airplanes:
    (1) The instruments required by FAR 23.1305 (a)(1) through (4), 
(b)(2) and (4).
    (2) A gas temperature indicator for each engine.
    (3) Free air temperature indicator.
    (4) A fuel flowmeter indicator for each engine.
    (5) Oil pressure warning means for each engine.
    (6) A torque indicator or adequate means for indicating power output 
for each engine.
    (7) Fire warning indicator for each engine.
    (8) A means to indicate when the propeller blade angle is below the 
low-pitch position corresponding to idle operation in flight.
    (9) A means to indicate the functioning of the ice protection system 
for each engine.
    (b) For turbopropeller powered airplanes, the turbopropeller blade 
position indicator must begin indicating when the blade has moved below 
the flight low-pitch position.
    (c) The following instruments are required for reciprocating-engine 
powered airplanes:
    (1) The instruments required by FAR 23.1305.
    (2) A cylinder head temperature indicator for each engine.
    (3) A manifold pressure indicator for each engine.

                         Systems and Equipments

                                 general

    59. Function and installation. The systems and equipment of the 
airplane must meet the requirements of FAR 23.1301, and the following:
    (a) Each item of additional installed equipment must--
    (1) Be of a kind and design appropriate to its intended function;
    (2) Be labeled as to its identification, function, or operating 
limitations, or any applicable combination of these factors, unless 
misuse or inadvertent actuation cannot create a hazard;
    (3) Be installed according to limitations specified for that 
equipment; and
    (4) Function properly when installed.
    (b) Systems and installations must be designed to safeguard against 
hazards to the aircraft in the event of their malfunction or failure.
    (c) Where an installation, the functioning of which is necessary in 
showing compliance with the applicable requirements, requires a power 
supply, such installation must be considered an essential load on the 
power supply, and the power sources and the distribution system must be 
capable of supplying the following power loads in probable operation 
combinations and for probable durations:
    (1) All essential loads after failure of any prime mover, power 
converter, or energy storage device.
    (2) All essential loads after failure of any one engine on two-
engine airplanes.
    (3) In determining the probable operating combinations and durations 
of essential loads for the power failure conditions described in 
subparagraphs (1) and (2) of this paragraph, it is permissible to assume 
that the power loads are reduced in accordance with a monitoring 
procedure which is consistent with safety in the types of operations 
authorized.
    60. Ventilation. The ventilation system of the airplane must meet 
the requirements of FAR 23.831, and in addition, for pressurized 
aircraft the ventilating air in flight crew and passenger compartments 
must be free of harmful or hazardous concentrations of gases and vapors 
in normal operation and in the event of reasonably probable failures or 
malfunctioning of the ventilating, heating, pressurization, or other 
systems, and equipment. If accumulation of hazardous quantities of smoke 
in the cockpit area is reasonably probable, smoke evacuation must be 
readily accomplished.

                    Electrical Systems and Equipment

    61. General. The electrical systems and equipment of the airplane 
must meet the requirements of FAR 23.1351, and the following:
    (a) Electrical system capacity. The required generating capacity, 
and number and kinds of power sources must--
    (1) Be determined by an electrical load analysis, and
    (2) Meet the requirements of FAR 23.1301.
    (b) Generating system. The generating system includes electrical 
power sources, main power busses, transmission cables, and associated 
control, regulation, and protective devices. It must be designed so 
that--
    (1) The system voltage and frequency (as applicable) at the 
terminals of all essential load equipment can be maintained within the 
limits for which the equipment is designed, during any probable 
operating conditions;
    (2) System transients due to switching, fault clearing, or other 
causes do not make essential loads inoperative, and do not cause a smoke 
or fire hazard;
    (3) There are means, accessible in flight to appropriate 
crewmembers, for the individual and collective disconnection of the 
electrical power sources from the system; and
    (4) There are means to indicate to appropriate crewmembers the 
generating system

[[Page 185]]

quantities essential for the safe operation of the system, including the 
voltage and current supplied by each generator.
    62. Electrical equipment and installation. Electrical equipment 
controls, and wiring must be installed so that operation of any one unit 
or system of units will not adversely affect the simultaneous operation 
of to the safe operation.
    63. Distribution system. (a) For the purpose of complying with this 
section, the distribution system includes the distribution busses, their 
associated feeders and each control and protective device.
    (b) Each system must be designed so that essential load circuits can 
be supplied in the event of reasonably probable faults or open circuits, 
including faults in heavy current carrying cables.
    (c) If two independent sources of electrical power for particular 
equipment or systems are required by this regulation, their electrical 
energy supply must be insured by means such as duplicate electrical 
equipment, throwover switching, or multichannel or loop circuits 
separately routed.
    64. Circuit protective devices. The circuit protective devices for 
the electrical circuits of the airplane must meet the requirements of 
FAR 23.1357, and in addition circuits for loads which are essential to 
safe operation must have individual and exclusive circuit protection.

[Doc. No. 8070, 34 FR 189, Jan. 7, 1969, as amended by SFAR 23-1, 34 FR 
20176, Dec. 24, 1969; 35 FR 1102, Jan. 28, 1970]



                            Subpart A_General



Sec. 23.1  Applicability.

    (a) This part prescribes airworthiness standards for the issue of 
type certificates, and changes to those certificates, for airplanes in 
the normal, utility, acrobatic, and commuter categories.
    (b) Each person who applies under Part 21 for such a certificate or 
change must show compliance with the applicable requirements of this 
part.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-34, 
52 FR 1825, Jan. 15, 1987]



Sec. 23.2  Special retroactive requirements.

    (a) Notwithstanding Sec. Sec. 21.17 and 21.101 of this chapter and 
irrespective of the type certification basis, each normal, utility, and 
acrobatic category airplane having a passenger seating configuration, 
excluding pilot seats, of nine or less, manufactured after December 12, 
1986, or any such foreign airplane for entry into the United States must 
provide a safety belt and shoulder harness for each forward- or aft-
facing seat which will protect the occupant from serious head injury 
when subjected to the inertia loads resulting from the ultimate static 
load factors prescribed in Sec. 23.561(b)(2) of this part, or which 
will provide the occupant protection specified in Sec. 23.562 of this 
part when that section is applicable to the airplane. For other seat 
orientations, the seat/restraint system must be designed to provide a 
level of occupant protection equivalent to that provided for forward- or 
aft-facing seats with a safety belt and shoulder harness installed.
    (b) Each shoulder harness installed at a flight crewmember station, 
as required by this section, must allow the crewmember, when seated with 
the safety belt and shoulder harness fastened, to perform all functions 
necessary for flight operations.
    (c) For the purpose of this section, the date of manufacture is:
    (1) The date the inspection acceptance records, or equivalent, 
reflect that the airplane is complete and meets the FAA approved type 
design data; or
    (2) In the case of a foreign manufactured airplane, the date the 
foreign civil airworthiness authority certifies the airplane is complete 
and issues an original standard airworthiness certificate, or the 
equivalent in that country.

[Amdt. 23-36, 53 FR 30812, Aug. 15, 1988]



Sec. 23.3  Airplane categories.

    (a) The normal category is limited to airplanes that have a seating 
configuration, excluding pilot seats, of nine or less, a maximum 
certificated takeoff weight of 12,500 pounds or less, and intended for 
nonacrobatic operation. Nonacrobatic operation includes:
    (1) Any maneuver incident to normal flying;
    (2) Stalls (except whip stalls); and
    (3) Lazy eights, chandelles, and steep turns, in which the angle of 
bank is not more than 60 degrees.
    (b) The utility category is limited to airplanes that have a seating 
configuration, excluding pilot seats, of nine or

[[Page 186]]

less, a maximum certificated takeoff weight of 12,500 pounds or less, 
and intended for limited acrobatic operation. Airplanes certificated in 
the utility category may be used in any of the operations covered under 
paragraph (a) of this section and in limited acrobatic operations. 
Limited acrobatic operation includes:
    (1) Spins (if approved for the particular type of airplane); and
    (2) Lazy eights, chandelles, and steep turns, or similar maneuvers, 
in which the angle of bank is more than 60 degrees but not more than 90 
degrees.
    (c) The acrobatic category is limited to airplanes that have a 
seating configuration, excluding pilot seats, of nine or less, a maximum 
certificated takeoff weight of 12,500 pounds or less, and intended for 
use without restrictions, other than those shown to be necessary as a 
result of required flight tests.
    (d) The commuter category is limited to multiengine airplanes that 
have a seating configuration, excluding pilot seats, of 19 or less, and 
a maximum certificated takeoff weight of 19,000 pounds or less. The 
commuter category operation is limited to any maneuver incident to 
normal flying, stalls (except whip stalls), and steep turns, in which 
the angle of bank is not more than 60 degrees.
    (e) Except for commuter category, airplanes may be type certificated 
in more than one category if the requirements of each requested category 
are met.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-4, 32 
FR 5934, Apr. 14, 1967; Amdt. 23-34, 52 FR 1825, Jan. 15, 1987; 52 FR 
34745, Sept. 14, 1987; Amdt. 23-50, 61 FR 5183, Feb. 9, 1996; Amdt. 23-
62, 76 FR 75753, Dec. 2, 2011]



                            Subpart B_Flight

                                 General



Sec. 23.21  Proof of compliance.

    (a) Each requirement of this subpart must be met at each appropriate 
combination of weight and center of gravity within the range of loading 
conditions for which certification is requested. This must be shown--
    (1) By tests upon an airplane of the type for which certification is 
requested, or by calculations based on, and equal in accuracy to, the 
results of testing; and
    (2) By systematic investigation of each probable combination of 
weight and center of gravity, if compliance cannot be reasonably 
inferred from combinations investigated.
    (b) The following general tolerances are allowed during flight 
testing. However, greater tolerances may be allowed in particular tests:

------------------------------------------------------------------------
                   Item                               Tolerance
------------------------------------------------------------------------
Weight....................................  +5%, -10%.
Critical items affected by weight.........  +5%, -1%.
C.G.......................................  7%
                                             total travel.
------------------------------------------------------------------------



Sec. 23.23  Load distribution limits.

    (a) Ranges of weights and centers of gravity within which the 
airplane may be safely operated must be established. If a weight and 
center of gravity combination is allowable only within certain lateral 
load distribution limits that could be inadvertently exceeded, these 
limits must be established for the corresponding weight and center of 
gravity combinations.
    (b) The load distribution limits may not exceed any of the 
following:
    (1) The selected limits;
    (2) The limits at which the structure is proven; or
    (3) The limits at which compliance with each applicable flight 
requirement of this subpart is shown.

[Doc. No. 26269, 58 FR 42156, Aug. 6, 1993]



Sec. 23.25  Weight limits.

    (a) Maximum weight. The maximum weight is the highest weight at 
which compliance with each applicable requirement of this part (other 
than those complied with at the design landing weight) is shown. The 
maximum weight must be established so that it is--
    (1) Not more than the least of--
    (i) The highest weight selected by the applicant; or
    (ii) The design maximum weight, which is the highest weight at which 
compliance with each applicable structural loading condition of this 
part (other than those complied with at the design landing weight) is 
shown; or

[[Page 187]]

    (iii) The highest weight at which compliance with each applicable 
flight requirement is shown, and
    (2) Not less than the weight with--
    (i) Each seat occupied, assuming a weight of 170 pounds for each 
occupant for normal and commuter category airplanes, and 190 pounds for 
utility and acrobatic category airplanes, except that seats other than 
pilot seats may be placarded for a lesser weight; and
    (A) Oil at full capacity, and
    (B) At least enough fuel for maximum continuous power operation of 
at least 30 minutes for day-VFR approved airplanes and at least 45 
minutes for night-VFR and IFR approved airplanes; or
    (ii) The required minimum crew, and fuel and oil to full tank 
capacity.
    (b) Minimum weight. The minimum weight (the lowest weight at which 
compliance with each applicable requirement of this part is shown) must 
be established so that it is not more than the sum of--
    (1) The empty weight determined under Sec. 23.29;
    (2) The weight of the required minimum crew (assuming a weight of 
170 pounds for each crewmember); and
    (3) The weight of--
    (i) For turbojet powered airplanes, 5 percent of the total fuel 
capacity of that particular fuel tank arrangement under investigation, 
and
    (ii) For other airplanes, the fuel necessary for one-half hour of 
operation at maximum continuous power.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13086, Aug. 13, 1969; Amdt. 23-21, 43 FR 2317, Jan. 16, 1978; Amdt. 
23-34, 52 FR 1825, Jan. 15, 1987; Amdt. 23-45, 58 FR 42156, Aug. 6, 
1993; Amdt. 23-50, 61 FR 5183, Feb. 9, 1996]



Sec. 23.29  Empty weight and corresponding center of gravity.

    (a) The empty weight and corresponding center of gravity must be 
determined by weighing the airplane with--
    (1) Fixed ballast;
    (2) Unusable fuel determined under Sec. 23.959; and
    (3) Full operating fluids, including--
    (i) Oil;
    (ii) Hydraulic fluid; and
    (iii) Other fluids required for normal operation of airplane 
systems, except potable water, lavatory precharge water, and water 
intended for injection in the engines.
    (b) The condition of the airplane at the time of determining empty 
weight must be one that is well defined and can be easily repeated.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-21, 43 FR 2317, Jan. 16, 1978]



Sec. 23.31  Removable ballast.

    Removable ballast may be used in showing compliance with the flight 
requirements of this subpart, if--
    (a) The place for carrying ballast is properly designed and 
installed, and is marked under Sec. 23.1557; and
    (b) Instructions are included in the airplane flight manual, 
approved manual material, or markings and placards, for the proper 
placement of the removable ballast under each loading condition for 
which removable ballast is necessary.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-13, 37 FR 20023, Sept. 23, 1972]



Sec. 23.33  Propeller speed and pitch limits.

    (a) General. The propeller speed and pitch must be limited to values 
that will assure safe operation under normal operating conditions.
    (b) Propellers not controllable in flight. For each propeller whose 
pitch cannot be controlled in flight--
    (1) During takeoff and initial climb at the all engine(s) operating 
climb speed specified in Sec. 23.65, the propeller must limit the 
engine r.p.m., at full throttle or at maximum allowable takeoff manifold 
pressure, to a speed not greater than the maximum allowable takeoff 
r.p.m.; and
    (2) During a closed throttle glide, at VNE, the propeller may not 
cause an engine speed above 110 percent of maximum continuous speed.
    (c) Controllable pitch propellers without constant speed controls. 
Each propeller that can be controlled in flight, but that does not have 
constant speed controls, must have a means to limit the pitch range so 
that--

[[Page 188]]

    (1) The lowest possible pitch allows compliance with paragraph 
(b)(1) of this section; and
    (2) The highest possible pitch allows compliance with paragraph 
(b)(2) of this section.
    (d) Controllable pitch propellers with constant speed controls. Each 
controllable pitch propeller with constant speed controls must have--
    (1) With the governor in operation, a means at the governor to limit 
the maximum engine speed to the maximum allowable takeoff r.p.m.; and
    (2) With the governor inoperative, the propeller blades at the 
lowest possible pitch, with takeoff power, the airplane stationary, and 
no wind, either--
    (i) A means to limit the maximum engine speed to 103 percent of the 
maximum allowable takeoff r.p.m., or
    (ii) For an engine with an approved overspeed, a means to limit the 
maximum engine and propeller speed to not more than the maximum approved 
overspeed.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42156, Aug. 6, 1993; Amdt. 23-50, 61 FR 5183, Feb. 9, 1996]

                               Performance



Sec. 23.45  General.

    (a) Unless otherwise prescribed, the performance requirements of 
this part must be met for--
    (1) Still air and standard atmosphere; and
    (2) Ambient atmospheric conditions, for commuter category airplanes, 
for reciprocating engine-powered airplanes of more than 6,000 pounds 
maximum weight, and for turbine engine-powered airplanes.
    (b) Performance data must be determined over not less than the 
following ranges of conditions--
    (1) Airport altitudes from sea level to 10,000 feet; and
    (2) For reciprocating engine-powered airplanes of 6,000 pounds, or 
less, maximum weight, temperature from standard to 30 [deg]C above 
standard; or
    (3) For reciprocating engine-powered airplanes of more than 6,000 
pounds maximum weight and turbine engine-powered airplanes, temperature 
from standard to 30 [deg]C above standard, or the maximum ambient 
atmospheric temperature at which compliance with the cooling provisions 
of Sec. 23.1041 to Sec. 23.1047 is shown, if lower.
    (c) Performance data must be determined with the cowl flaps or other 
means for controlling the engine cooling air supply in the position used 
in the cooling tests required by Sec. Sec. 23.1041 to 23.1047.
    (d) The available propulsive thrust must correspond to engine power, 
not exceeding the approved power, less--
    (1) Installation losses; and
    (2) The power absorbed by the accessories and services appropriate 
to the particular ambient atmospheric conditions and the particular 
flight condition.
    (e) The performance, as affected by engine power or thrust, must be 
based on a relative humidity:
    (1) Of 80 percent at and below standard temperature; and
    (2) From 80 percent, at the standard temperature, varying linearly 
down to 34 percent at the standard temperature plus 50 [deg]F.
    (f) Unless otherwise prescribed, in determining the takeoff and 
landing distances, changes in the airplane's configuration, speed, and 
power must be made in accordance with procedures established by the 
applicant for operation in service. These procedures must be able to be 
executed consistently by pilots of average skill in atmospheric 
conditions reasonably expected to be encountered in service.
    (g) The following, as applicable, must be determined on a smooth, 
dry, hard-surfaced runway--
    (1) Takeoff distance of Sec. 23.53(b);
    (2) Accelerate-stop distance of Sec. 23.55;
    (3) Takeoff distance and takeoff run of Sec. 23.59; and
    (4) Landing distance of Sec. 23.75.

    Note: The effect on these distances of operation on other types of 
surfaces (for example, grass, gravel) when dry, may be determined or 
derived and these surfaces listed in the Airplane Flight Manual in 
accordance with Sec. 23.1583(p).

    (h) For multiengine jets weighing over 6,000 pounds in the normal, 
utility, and acrobatic category and commuter category airplanes, the 
following also apply:
    (1) Unless otherwise prescribed, the applicant must select the 
takeoff,

[[Page 189]]

enroute, approach, and landing configurations for the airplane.
    (2) The airplane configuration may vary with weight, altitude, and 
temperature, to the extent that they are compatible with the operating 
procedures required by paragraph (h)(3) of this section.
    (3) Unless otherwise prescribed, in determining the critical-engine-
inoperative takeoff performance, takeoff flight path, and accelerate-
stop distance, changes in the airplane's configuration, speed, and power 
must be made in accordance with procedures established by the applicant 
for operation in service.
    (4) Procedures for the execution of discontinued approaches and 
balked landings associated with the conditions prescribed in Sec. 
23.67(c)(4) and Sec. 23.77(c) must be established.
    (5) The procedures established under paragraphs (h)(3) and (h)(4) of 
this section must--
    (i) Be able to be consistently executed by a crew of average skill 
in atmospheric conditions reasonably expected to be encountered in 
service;
    (ii) Use methods or devices that are safe and reliable; and
    (iii) Include allowance for any reasonably expected time delays in 
the execution of the procedures.

[Doc. No. 27807, 61 FR 5184, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75753, Dec. 2, 2011]



Sec. 23.49  Stalling speed.

    (a) VSO (maximum landing flap configuration) and 
VS1 are the stalling speeds or the minimum steady flight 
speeds, in knots (CAS), at which the airplane is controllable with--
    (1) For reciprocating engine-powered airplanes, the engine(s) 
idling, the throttle(s) closed or at not more than the power necessary 
for zero thrust at a speed not more than 110 percent of the stalling 
speed;
    (2) For turbine engine-powered airplanes, the propulsive thrust not 
greater than zero at the stalling speed, or, if the resultant thrust has 
no appreciable effect on the stalling speed, with engine(s) idling and 
throttle(s) closed;
    (3) The propeller(s) in the takeoff position;
    (4) The airplane in the condition existing in the test, in which 
VSO and VS1 are being used;
    (5) The center of gravity in the position that results in the 
highest value of VSO and VS1; and
    (6) The weight used when VSO and VS1 are being 
used as a factor to determine compliance with a required performance 
standard.
    (b) VSO and VS1 must be determined by flight 
tests, using the procedure and meeting the flight characteristics 
specified in Sec. 23.201.
    (c) Except as provided in paragraph (d) of this section, 
VSO at maximum weight may not exceed 61 knots for--
    (1) Single-engine airplanes; and
    (2) Multiengine airplanes of 6,000 pounds or less maximum weight 
that cannot meet the minimum rate of climb specified in Sec. 23.67(a) 
(1) with the critical engine inoperative.
    (d) All single-engine airplanes, and those multiengine airplanes of 
6,000 pounds or less maximum weight with a VSO of more than 
61 knots that do not meet the requirements of Sec. 23.67(a)(1), must 
comply with Sec. 23.562(d).

[Doc. No. 27807, 61 FR 5184, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75753, Dec. 2, 2011]



Sec. 23.51  Takeoff speeds.

    (a) For normal, utility, and acrobatic category airplanes, rotation 
speed, VR, is the speed at which the pilot makes a control 
input, with the intention of lifting the airplane out of contact with 
the runway or water surface.
    (1) For multiengine landplanes, VR, must not be less than 
the greater of 1.05 VMC; or 1.10 VS1;
    (2) For single-engine landplanes, VR, must not be less 
than VS1; and
    (3) For seaplanes and amphibians taking off from water, 
VR, may be any speed that is shown to be safe under all 
reasonably expected conditions, including turbulence and complete 
failure of the critical engine.
    (b) For normal, utility, and acrobatic category airplanes, the speed 
at 50 feet above the takeoff surface level must not be less than:
    (1) For multiengine airplanes, the highest of--
    (i) A speed that is shown to be safe for continued flight (or 
emergency

[[Page 190]]

landing, if applicable) under all reasonably expected conditions, 
including turbulence and complete failure of the critical engine;
    (ii) 1.10 VMC; or
    (iii) 1.20 VS1.
    (2) For single-engine airplanes, the higher of--
    (i) A speed that is shown to be safe under all reasonably expected 
conditions, including turbulence and complete engine failure; or
    (ii) 1.20 VS1.
    (c) For normal, utility, and acrobatic category multiengine jets of 
more than 6,000 pounds maximum weight and commuter category airplanes, 
the following apply:
    (l) V1 must be established in relation to VEF 
as follows:
    (i) VEF is the calibrated airspeed at which the critical 
engine is assumed to fail. VEF must be selected by the 
applicant but must not be less than 1.05 VMC determined under 
Sec. 23.149(b) or, at the option of the applicant, not less than 
VMCG determined under Sec. 23.149(f).
    (ii) The takeoff decision speed, V1, is the calibrated 
airspeed on the ground at which, as a result of engine failure or other 
reasons, the pilot is assumed to have made a decision to continue or 
discontinue the takeoff. The takeoff decision speed, V1, must 
be selected by the applicant but must not be less than VEF 
plus the speed gained with the critical engine inoperative during the 
time interval between the instant at which the critical engine is failed 
and the instant at which the pilot recognizes and reacts to the engine 
failure, as indicated by the pilot's application of the first retarding 
means during the accelerate-stop determination of Sec. 23.55.
    (2) The rotation speed, VR, in terms of calibrated 
airspeed, must be selected by the applicant and must not be less than 
the greatest of the following:
    (i) V1;
    (ii) 1.05 VMC determined under Sec. 23.149(b);
    (iii) 1.10 VS1; or
    (iv) The speed that allows attaining the initial climb-out speed, 
V2, before reaching a height of 35 feet above the takeoff 
surface in accordance with Sec. 23.57(c)(2).
    (3) For any given set of conditions, such as weight, altitude, 
temperature, and configuration, a single value of VR must be 
used to show compliance with both the one-engine-inoperative takeoff and 
all-engines-operating takeoff requirements.
    (4) The takeoff safety speed, V2, in terms of calibrated 
airspeed, must be selected by the applicant so as to allow the gradient 
of climb required in Sec. 23.67 (c)(1) and (c)(2) but mut not be less 
than 1.10 VMC or less than 1.20 VS1.
    (5) The one-engine-inoperative takeoff distance, using a normal 
rotation rate at a speed 5 knots less than VR, established in 
accordance with paragraph (c)(2) of this section, must be shown not to 
exceed the corresponding one-engine-inoperative takeoff distance, 
determined in accordance with Sec. 23.57 and Sec. 23.59(a)(1), using 
the established VR. The takeoff, otherwise performed in 
accordance with Sec. 23.57, must be continued safely from the point at 
which the airplane is 35 feet above the takeoff surface and at a speed 
not less than the established V2 minus 5 knots.
    (6) The applicant must show, with all engines operating, that marked 
increases in the scheduled takeoff distances, determined in accordance 
with Sec. 23.59(a)(2), do not result from over-rotation of the airplane 
or out-of-trim conditions.

[Doc. No. 27807, 61 FR 5184, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75753, Dec. 2, 2011]



Sec. 23.53  Takeoff performance.

    (a) For normal, utility, and acrobatic category airplanes, the 
takeoff distance must be determined in accordance with paragraph (b) of 
this section, using speeds determined in accordance with Sec. 23.51 (a) 
and (b).
    (b) For normal, utility, and acrobatic category airplanes, the 
distance required to takeoff and climb to a height of 50 feet above the 
takeoff surface must be determined for each weight, altitude, and 
temperature within the operational limits established for takeoff with--
    (1) Takeoff power on each engine;
    (2) Wing flaps in the takeoff position(s); and
    (3) Landing gear extended.
    (c) For normal, utility, and acrobatic category multiengine jets of 
more than 6,000 pounds maximum weight and

[[Page 191]]

commuter category airplanes, takeoff performance, as required by 
Sec. Sec. 23.55 through 23.59, must be determined with the operating 
engine(s) within approved operating limitations.

[Doc. No. 27807, 61 FR 5185, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75753, Dec. 2, 2011]



Sec. 23.55  Accelerate-stop distance.

    For normal, utility, and acrobatic category multiengine jets of more 
than 6,000 pounds maximum weight and commuter category airplanes, the 
accelerate-stop distance must be determined as follows:
    (a) The accelerate-stop distance is the sum of the distances 
necessary to--
    (1) Accelerate the airplane from a standing start to VEF 
with all engines operating;
    (2) Accelerate the airplane from VEF to V1, 
assuming the critical engine fails at VEF; and
    (3) Come to a full stop from the point at which V1 is 
reached.
    (b) Means other than wheel brakes may be used to determine the 
accelerate-stop distances if that means--
    (1) Is safe and reliable;
    (2) Is used so that consistent results can be expected under normal 
operating conditions; and
    (3) Is such that exceptional skill is not required to control the 
airplane.

[Amdt. 23-34, 52 FR 1826, Jan. 15, 1987, as amended by Amdt. 23-50, 61 
FR 5185, Feb. 9, 1996, as amended by Amdt. 23-62, 76 FR 75753, Dec. 2, 
2011]



Sec. 23.57  Takeoff path.

    For normal, utility, and acrobatic category multiengine jets of more 
than 6,000 pounds maximum weight and commuter category airplanes, the 
takeoff path is as follows:
    (a) The takeoff path extends from a standing start to a point in the 
takeoff at which the airplane is 1500 feet above the takeoff surface at 
or below which height the transition from the takeoff to the enroute 
configuration must be completed; and
    (1) The takeoff path must be based on the procedures prescribed in 
Sec. 23.45;
    (2) The airplane must be accelerated on the ground to VEF at which 
point the critical engine must be made inoperative and remain 
inoperative for the rest of the takeoff; and
    (3) After reaching VEF, the airplane must be accelerated to V2.
    (b) During the acceleration to speed V2, the nose gear 
may be raised off the ground at a speed not less than VR. 
However, landing gear retraction must not be initiated until the 
airplane is airborne.
    (c) During the takeoff path determination, in accordance with 
paragraphs (a) and (b) of this section--
    (1) The slope of the airborne part of the takeoff path must not be 
negative at any point;
    (2) The airplane must reach V2 before it is 35 feet above the 
takeoff surface, and must continue at a speed as close as practical to, 
but not less than V2, until it is 400 feet above the takeoff surface;
    (3) At each point along the takeoff path, starting at the point at 
which the airplane reaches 400 feet above the takeoff surface, the 
available gradient of climb must not be less than--
    (i) 1.2 percent for two-engine airplanes;
    (ii) 1.5 percent for three-engine airplanes;
    (iii) 1.7 percent for four-engine airplanes; and
    (4) Except for gear retraction and automatic propeller feathering, 
the airplane configuration must not be changed, and no change in power 
that requires action by the pilot may be made, until the airplane is 400 
feet above the takeoff surface.
    (d) The takeoff path to 35 feet above the takeoff surface must be 
determined by a continuous demonstrated takeoff.
    (e) The takeoff path to 35 feet above the takeoff surface must be 
determined by synthesis from segments; and
    (1) The segments must be clearly defined and must be related to 
distinct changes in configuration, power, and speed;
    (2) The weight of the airplane, the configuration, and the power 
must be assumed constant throughout each segment and must correspond to 
the most critical condition prevailing in the segment; and

[[Page 192]]

    (3) The takeoff flight path must be based on the airplane's 
performance without utilizing ground effect.

[Amdt. 23-34, 52 FR 1827, Jan. 15, 1987, as amended by Amdt. 23-50, 61 
FR 5185, Feb. 9, 1996, as amended by Amdt. 23-62, 76 FR 75753, Dec. 2, 
2011]



Sec. 23.59  Takeoff distance and takeoff run.

    For normal, utility, and acrobatic category multiengine jets of more 
than 6,000 pounds maximum weight and commuter category airplanes, the 
takeoff distance and, at the option of the applicant, the takeoff run, 
must be determined.
    (a) Takeoff distance is the greater of--
    (1) The horizontal distance along the takeoff path from the start of 
the takeoff to the point at which the airplane is 35 feet above the 
takeoff surface as determined under Sec. 23.57; or
    (2) With all engines operating, 115 percent of the horizontal 
distance from the start of the takeoff to the point at which the 
airplane is 35 feet above the takeoff surface, determined by a procedure 
consistent with Sec. 23.57.
    (b) If the takeoff distance includes a clearway, the takeoff run is 
the greater of--
    (1) The horizontal distance along the takeoff path from the start of 
the takeoff to a point equidistant between the liftoff point and the 
point at which the airplane is 35 feet above the takeoff surface as 
determined under Sec. 23.57; or
    (2) With all engines operating, 115 percent of the horizontal 
distance from the start of the takeoff to a point equidistant between 
the liftoff point and the point at which the airplane is 35 feet above 
the takeoff surface, determined by a procedure consistent with Sec. 
23.57.

[Amdt. 23-34, 52 FR 1827, Jan. 15, 1987, as amended by Amdt. 23-50, 61 
FR 5185, Feb. 9, 1996, as amended by Amdt. 23-62, 76 FR 75753, Dec. 2, 
2011]



Sec. 23.61  Takeoff flight path.

    For normal, utility, and acrobatic category multiengine jets of more 
than 6,000 pounds maximum weight and commuter category airplanes, the 
takeoff flight path must be determined as follows:
    (a) The takeoff flight path begins 35 feet above the takeoff surface 
at the end of the takeoff distance determined in accordance with Sec. 
23.59.
    (b) The net takeoff flight path data must be determined so that they 
represent the actual takeoff flight paths, as determined in accordance 
with Sec. 23.57 and with paragraph (a) of this section, reduced at each 
point by a gradient of climb equal to--
    (1) 0.8 percent for two-engine airplanes;
    (2) 0.9 percent for three-engine airplanes; and
    (3) 1.0 percent for four-engine airplanes.
    (c) The prescribed reduction in climb gradient may be applied as an 
equivalent reduction in acceleration along that part of the takeoff 
flight path at which the airplane is accelerated in level flight.

[Amdt. 23-34, 52 FR 1827, Jan. 15, 1987, as amended by Amdt. 23-62, 76 
FR 75753, Dec. 2, 2011]



Sec. 23.63  Climb: General.

    (a) Compliance with the requirements of Sec. Sec. 23.65, 23.66, 
23.67, 23.69, and 23.77 must be shown--
    (1) Out of ground effect; and
    (2) At speeds that are not less than those at which compliance with 
the powerplant cooling requirements of Sec. Sec. 23.1041 to 23.1047 has 
been demonstrated; and
    (3) Unless otherwise specified, with one engine inoperative, at a 
bank angle not exceeding 5 degrees.
    (b) For normal, utility, and acrobatic category reciprocating 
engine-powered airplanes of 6,000 pounds or less maximum weight, 
compliance must be shown with Sec. 23.65(a), Sec. 23.67(a), where 
appropriate, and Sec. 23.77(a) at maximum takeoff or landing weight, as 
appropriate, in a standard atmosphere.
    (c) For reciprocating engine-powered airplanes of more than 6,000 
pounds maximum weight, single-engine turbines, and multiengine turbine 
airplanes of 6,000 pounds or less maximum weight in the normal, utility, 
and acrobatic category, compliance must be shown at weights as a 
function of airport altitude and ambient temperature,

[[Page 193]]

within the operational limits established for takeoff and landing, 
respectively, with--
    (1) Sections 23.65(b) and 23.67(b) (1) and (2), where appropriate, 
for takeoff, and
    (2) Section 23.67(b)(2), where appropriate, and Sec. 23.77(b), for 
landing.
    (d) For multiengine turbine airplanes over 6,000 pounds maximum 
weight in the normal, utility, and acrobatic category and commuter 
category airplanes, compliance must be shown at weights as a function of 
airport altitude and ambient temperature within the operational limits 
established for takeoff and landing, respectively, with--
    (1) Sections 23.67(c)(1), 23.67(c)(2), and 23.67(c)(3) for takeoff; 
and
    (2) Sections 23.67(c)(3), 23.67(c)(4), and 23.77(c) for landing.

[Doc. No. 27807, 61 FR 5186, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75753, Dec. 2, 2011]



Sec. 23.65  Climb: All engines operating.

    (a) Each normal, utility, and acrobatic category reciprocating 
engine-powered airplane of 6,000 pounds or less maximum weight must have 
a steady climb gradient at sea level of at least 8.3 percent for 
landplanes or 6.7 percet for seaplanes and amphibians with--
    (1) Not more than maximum continuous power on each engine;
    (2) The landing gear retracted;
    (3) The wing flaps in the takeoff position(s); and
    (4) A climb speed not less than the greater of 1.1 VMC 
and 1.2 VS1 for multiengine airplanes and not less than 1.2 
VS1 for single--engine airplanes.
    (b) Each normal, utility, and acrobatic category reciprocating 
engine-powered airplane of more than 6,000 pounds maximum weight, 
single-engine turbine, and multiengine turbine airplanes of 6,000 pounds 
or less maximum weight in the normal, utility, and acrobatic category 
must have a steady gradient of climb after takeoff of at least 4 percent 
with
    (1) Take off power on each engine;
    (2) The landing gear extended, except that if the landing gear can 
be retracted in not more than seven seconds, the test may be conducted 
with the gear retracted;
    (3) The wing flaps in the takeoff position(s); and
    (4) A climb speed as specified in Sec. 23.65(a)(4).

[Doc. No. 27807, 61 FR 5186, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75753, Dec. 2, 2011]



Sec. 23.66  Takeoff climb: One-engine inoperative.

    For normal, utility, and acrobatic category reciprocating engine-
powered airplanes of more than 6,000 pounds maximum weight, and turbine 
engine-powered airplanes in the normal, utility, and acrobatic category, 
the steady gradient of climb or descent must be determined at each 
weight, altitude, and ambient temperature within the operational limits 
established by the applicant with--
    (a) The critical engine inoperative and its propeller in the 
position it rapidly and automatically assumes;
    (b) The remaining engine(s) at takeoff power;
    (c) The landing gear extended, except that if the landing gear can 
be retracted in not more than seven seconds, the test may be conducted 
with the gear retracted;
    (d) The wing flaps in the takeoff position(s):
    (e) The wings level; and
    (f) A climb speed equal to that achieved at 50 feet in the 
demonstration of Sec. 23.53.

[Doc. No. 27807, 61 FR 5186, Feb. 9, 1996]



Sec. 23.67  Climb: One engine inoperative.

    (a) For normal, utility, and acrobatic category reciprocating 
engine-powered airplanes of 6,000 pounds or less maximum weight, the 
following apply:
    (1) Except for those airplanes that meet the requirements prescribed 
in Sec. 23.562(d), each airplane with a VSO of more than 61 
knots must be able to maintain a steady climb gradient of at least 1.5 
percent at a pressure altitude of 5,000 feet with the--
    (i) Critical engine inoperative and its propeller in the minimum 
drag position;
    (ii) Remaining engine(s) at not more than maximum continuous power;
    (iii) Landing gear retracted;
    (iv) Wing flaps retracted; and
    (v) Climb speed not less than 1.2 VS1.

[[Page 194]]

    (2) For each airplane that meets the requirements prescribed in 
Sec. 23.562(d), or that has a VSO of 61 knots or less, the 
steady gradient of climb or descent at a pressure altitude of 5,000 feet 
must be determined with the--
    (i) Critical engine inoperative and its propeller in the minimum 
drag position;
    (ii) Remaining engine(s) at not more than maximum continuous power;
    (iii) Landing gear retracted;
    (iv) Wing flaps retracted; and
    (v) Climb speed not less than 1.2VS1.
    (b) For normal, utility, and acrobatic category reciprocating 
engine-powered airplanes of more than 6,000 pounds maximum weight, and 
turbopropeller-powered airplanes in the normal, utility, and acrobatic 
category--
    (1) The steady gradient of climb at an altitude of 400 feet above 
the takeoff must be no less than 1 percent with the--
    (i) Critical engine inoperative and its propeller in the minimum 
drag position;
    (ii) Remaining engine(s) at takeoff power;
    (iii) Landing gear retracted;
    (iv) Wing flaps in the takeoff position(s); and
    (v) Climb speed equal to that achieved at 50 feet in the 
demonstration of Sec. 23.53.
    (2) The steady gradient of climb must not be less than 0.75 percent 
at an altitude of 1,500 feet above the takeoff surface, or landing 
surface, as appropriate, with the--
    (i) Critical engine inoperative and its propeller in the minimum 
drag position;
    (ii) Remaining engine(s) at not more than maximum continuous power;
    (iii) Landing gear retracted;
    (iv) Wing flaps retracted; and
    (v) Climb speed not less than 1.2 VS1.
    (c) For normal, utility, and acrobatic category jets of 6,000 pounds 
or less maximum weight--
    (1) The steady gradient of climb at an altitude of 400 feet above 
the takeoff must be no less than 1.2 percent with the--
    (i) Critical engine inoperative;
    (ii) Remaining engine(s) at takeoff power;
    (iii) Landing gear retracted;
    (iv) Wing flaps in the takeoff position(s); and
    (v) Climb speed equal to that achieved at 50 feet in the 
demonstration of Sec. 23.53.
    (2) The steady gradient of climb may not be less than 0.75 percent 
at an altitude of 1,500 feet above the takeoff surface, or landing 
surface, as appropriate, with the--
    (i) Critical engine inoperative;
    (ii) Remaining engine(s) at not more than maximum continuous power;
    (iii) Landing gear retracted;
    (iv) Wing flaps retracted; and
    (v) Climb speed not less than 1.2 VS1.
    (d) For jets over 6,000 pounds maximum weight in the normal, utility 
and acrobatic category and commuter category airplanes, the following 
apply:
    (1) Takeoff; landing gear extended. The steady gradient of climb at 
the altitude of the takeoff surface must be measurably positive for two-
engine airplanes, not less than 0.3 percent for three-engine airplanes, 
or 0.5 percent for four-engine airplanes with--
    (i) The critical engine inoperative and its propeller in the 
position it rapidly and automatically assumes;
    (ii) The remaining engine(s) at takeoff power;
    (iii) The landing gear extended, and all landing gear doors open;
    (iv) The wing flaps in the takeoff position(s);
    (v) The wings level; and
    (vi) A climb speed equal to V2.
    (2) Takeoff; landing gear retracted. The steady gradient of climb at 
an altitude of 400 feet above the takeoff surface must be not less than 
2.0 percent of two-engine airplanes, 2.3 percent for three-engine 
airplanes, and 2.6 percent for four-engine airplanes with--
    (i) The critical engine inoperative and its propeller in the 
position it rapidly and automatically assumes;
    (ii) The remaining engine(s) at takeoff power;
    (iii) The landing gear retracted;
    (iv) The wing flaps in the takeoff position(s);
    (v) A climb speed equal to V2.
    (3) Enroute. The steady gradient of climb at an altitude of 1,500 
feet above the takeoff or landing surface, as appropriate, must be not 
less than 1.2 percent for two-engine airplanes, 1.5

[[Page 195]]

percent for three-engine airplanes, and 1.7 percent for four-engine 
airplanes with--
    (i) The critical engine inoperative and its propeller in the minimum 
drag position;
    (ii) The remaining engine(s) at not more than maximum continuous 
power;
    (iii) The landing gear retracted;
    (iv) The wing flaps retracted; and
    (v) A climb speed not less than 1.2 VS1.
    (4) Discontinued approach. The steady gradient of climb at an 
altitude of 400 feet above the landing surface must be not less than 2.1 
percent for two-engine airplanes, 2.4 percent for three-engine 
airplanes, and 2.7 percent for four-engine airplanes, with--
    (i) The critical engine inoperative and its propeller in the minimum 
drag position;
    (ii) The remaining engine(s) at takeoff power;
    (iii) Landing gear retracted;
    (iv) Wing flaps in the approach position(s) in which VS1 
for these position(s) does not exceed 110 percent of the VS1 
for the related all-engines-operated landing position(s); and
    (v) A climb speed established in connection with normal landing 
procedures but not exceeding 1.5 VS1.

[Doc. No. 27807, 61 FR 5186, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75754, Dec. 2, 2011]



Sec. 23.69  Enroute climb/descent.

    (a) All engines operating. The steady gradient and rate of climb 
must be determined at each weight, altitude, and ambient temperature 
within the operational limits established by the applicant with--
    (1) Not more than maximum continuous power on each engine;
    (2) The landing gear retracted;
    (3) The wing flaps retracted; and
    (4) A climb speed not less than 1.3 VS1.
    (b) One engine inoperative. The steady gradient and rate of climb/
descent must be determined at each weight, altitude, and ambient 
temperature within the operational limits established by the applicant 
with--
    (1) The critical engine inoperative and its propeller in the minimum 
drag position;
    (2) The remaining engine(s) at not more than maximum continuous 
power;
    (3) The landing gear retracted;
    (4) The wing flaps retracted; and
    (5) A climb speed not less than 1.2 VS1.

[Doc. No. 27807, 61 FR 5187, Feb. 9, 1996]



Sec. 23.71  Glide: Single-engine airplanes.

    The maximum horizontal distance traveled in still air, in nautical 
miles, per 1,000 feet of altitude lost in a glide, and the speed 
necessary to achieve this must be determined with the engine 
inoperative, its propeller in the minimum drag position, and landing 
gear and wing flaps in the most favorable available position.

[Doc. No. 27807, 61 FR 5187, Feb. 9, 1996]



Sec. 23.73  Reference landing approach speed.

    (a) For normal, utility, and acrobatic category reciprocating 
engine-powered airplanes of 6,000 pounds or less maximum weight, the 
reference landing approach speed, VREF, may not be less than 
the greater of VMC, determined in Sec. 23.149(b) with the 
wing flaps in the most extended takeoff position, and 1.3 
VS1.
    (b) For normal, utility, and acrobatic category turbine powered 
airplanes of 6,000 pounds or less maximum weight, turboprops of more 
than 6,000 pounds maximum weight, and reciprocating engine-powered 
airplanes of more than 6,000 pounds maximum weight, the reference 
landing approach speed, VREF, may not be less than the 
greater of VMC, determined in Sec. 23.149(c), and 1.3 
VS1.
    (c) For normal, utility, and acrobatic category jets of more than 
6,000 pounds maximum weight and commuter category airplanes, the 
reference landing approach speed, VREF, may not be less than 
the greater of 1.05 VMC, determined in Sec. 23.149(c), and 
1.3 VS1.

[Amdt. 23-62, 76 FR 75754, Dec. 2, 2011]



Sec. 23.75  Landing distance.

    The horizontal distance necessary to land and come to a complete 
stop from

[[Page 196]]

a point 50 feet above the landing surface must be determined, for 
standard temperatures at each weight and altitude within the operational 
limits established for landing, as follows:
    (a) A steady approach at not less than VREF, determined 
in accordance with Sec. 23.73 (a), (b), or (c), as appropriate, must be 
maintained down to the 50 foot height and--
    (1) The steady approach must be at a gradient of descent not greater 
than 5.2 percent (3 degrees) down to the 50-foot height.
    (2) In addition, an applicant may demonstrate by tests that a 
maximum steady approach gradient steeper than 5.2 percent, down to the 
50-foot height, is safe. The gradient must be established as an 
operating limitation and the information necessary to display the 
gradient must be available to the pilot by an appropriate instrument.
    (b) A constant configuration must be maintained throughout the 
maneuver.
    (c) The landing must be made without excessive vertical acceleration 
or tendency to bounce, nose over, ground loop, porpoise, or water loop.
    (d) It must be shown that a safe transition to the balked landing 
conditions of Sec. 23.77 can be made from the conditions that exist at 
the 50 foot height, at maximum landing weight, or at the maximum landing 
weight for altitude and temperature of Sec. 23.63 (c)(2) or (d)(2), as 
appropriate.
    (e) The brakes must be used so as to not cause excessive wear of 
brakes or tires.
    (f) Retardation means other than wheel brakes may be used if that 
means--
    (1) Is safe and reliable; and
    (2) Is used so that consistent results can be expected in service.
    (g) If any device is used that depends on the operation of any 
engine, and the landing distance would be increased when a landing is 
made with that engine inoperative, the landing distance must be 
determined with that engine inoperative unless the use of other 
compensating means will result in a landing distance not more than that 
with each engine operating.

[Amdt. 23-21, 43 FR 2318, Jan. 16, 1978, as amended by Amdt. 23-34, 52 
FR 1828, Jan. 15, 1987; Amdt. 23-42, 56 FR 351, Jan. 3, 1991; Amdt. 23-
50, 61 FR 5187, Feb. 9, 1996]



Sec. 23.77  Balked landing.

    (a) Each normal, utility, and acrobatic category reciprocating 
engine-powered airplane at 6,000 pounds or less maximum weight must be 
able to maintain a steady gradient of climb at sea level of at least 3.3 
percent with--
    (1) Takeoff power on each engine;
    (2) The landing gear extended;
    (3) The wing flaps in the landing position, except that if the flaps 
may safely be retracted in two seconds or less without loss of altitude 
and without sudden changes of angle of attack, they may be retracted; 
and
    (4) A climb speed equal to VREF, as defined in Sec. 
23.73(a).
    (b) Each normal, utility, and acrobatic category reciprocating 
engine-powered and single engine turbine powered airplane of more than 
6,000 pounds maximum weight, and multiengine turbine engine-powered 
airplane of 6,000 pounds or less maximum weight in the normal, utility, 
and acrobatic category must be able to maintain a steady gradient of 
climb of at least 2.5 percent with--
    (1) Not more than the power that is available on each engine eight 
seconds after initiation of movement of the power controls from minimum 
flight-idle position;
    (2) The landing gear extended;
    (3) The wing flaps in the landing position; and
    (4) A climb speed equal to VREF, as defined in Sec. 
23.73(b).
    (c) Each normal, utility, and acrobatic multiengine turbine powered 
airplane over 6,000 pounds maximum weight and each commuter category 
airplane must be able to maintain a steady gradient of climb of at least 
3.2 percent with--
    (1) Not more than the power that is available on each engine eight 
seconds after initiation of movement of the power controls from the 
minimum flight idle position;
    (2) Landing gear extended;

[[Page 197]]

    (3) Wing flaps in the landing position; and
    (4) A climb speed equal to VREF, as defined in Sec. 
23.73(c).

[Doc. No. 27807, 61 FR 5187, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75754, Dec. 2, 2011]

                         Flight Characteristics



Sec. 23.141  General.

    The airplane must meet the requirements of Sec. Sec. 23.143 through 
23.253 at all practical loading conditions and operating altitudes for 
which certification has been requested, not exceeding the maximum 
operating altitude established under Sec. 23.1527, and without 
requiring exceptional piloting skill, alertness, or strength.

[Doc. No. 26269, 58 FR 42156, Aug. 6, 1993]

                   Controllability and Maneuverability



Sec. 23.143  General.

    (a) The airplane must be safely controllable and maneuverable during 
all flight phases including--
    (1) Takeoff;
    (2) Climb;
    (3) Level flight;
    (4) Descent;
    (5) Go-around; and
    (6) Landing (power on and power off) with the wing flaps extended 
and retracted.
    (b) It must be possible to make a smooth transition from one flight 
condition to another (including turns and slips) without danger of 
exceeding the limit load factor, under any probable operating condition 
(including, for multiengine airplanes, those conditions normally 
encountered in the sudden failure of any engine).
    (c) If marginal conditions exist with regard to required pilot 
strength, the control forces necessary must be determined by 
quantitative tests. In no case may the control forces under the 
conditions specified in paragraphs (a) and (b) of this section exceed 
those prescribed in the following table:

------------------------------------------------------------------------
    Values in pounds force applied to the
               relevant control                 Pitch     Roll     Yaw
------------------------------------------------------------------------
(a) For temporary application:
  Stick......................................       60       30  .......
  Wheel (Two hands on rim)...................       75       50  .......
  Wheel (One hand on rim)....................       50       25  .......
  Rudder Pedal...............................  .......  .......      150
(b) For prolonged application................       10        5       20
------------------------------------------------------------------------


[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31819, Nov. 19, 1973; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; 
Amdt. 23-45, 58 FR 42156, Aug. 6, 1993; Amdt. 23-50, 61 FR 5188, Feb. 9, 
1996]



Sec. 23.145  Longitudinal control.

    (a) With the airplane as nearly as possible in trim at 1.3 
VS1, it must be possible, at speeds below the trim speed, to 
pitch the nose downward so that the rate of increase in airspeed allows 
prompt acceleration to the trim speed with--
    (1) Maximum continuous power on each engine;
    (2) Power off; and
    (3) Wing flap and landing gear--
    (i) retracted, and
    (ii) extended.
    (b) Unless otherwise required, it must be possible to carry out the 
following maneuvers without requiring the application of single-handed 
control forces exceeding those specified in Sec. 23.143(c). The 
trimming controls must not be adjusted during the maneuvers:
    (1) With the landing gear extended, the flaps retracted, and the 
airplanes as nearly as possible in trim at 1.4 VS1, extend 
the flaps as rapidly as possible and allow the airspeed to transition 
from 1.4VS1 to 1.4 VSO:
    (i) With power off; and
    (ii) With the power necessary to maintain level flight in the 
initial condition.
    (2) With landing gear and flaps extended, power off, and the 
airplane as nearly as possible in trim at 1.3 VSO: quickly 
apply takeoff power and retract the flaps as rapidly as possible to the 
recommended go around setting and allow the airspeed to transition from 
1.3 VSO to 1.3 VS1. Retract the gear when a 
positive rate of climb is established.
    (3) With landing gear and flaps extended, in level flight, power 
necessary to attain level flight at 1.1 VSO, and the airplane 
as nearly as possible in trim, it must be possible to maintain 
approximately level flight while retracting the flaps as rapidly as 
possible with

[[Page 198]]

simultaneous application of not more than maximum continuous power. If 
gated flat positions are provided, the flap retraction may be 
demonstrated in stages with power and trim reset for level flight at 1.1 
VS1, in the initial configuration for each stage--
    (i) From the fully extended position to the most extended gated 
position;
    (ii) Between intermediate gated positions, if applicable; and
    (iii) From the least extended gated position to the fully retracted 
position.
    (4) With power off, flaps and landing gear retracted and the 
airplane as nearly as possible in trim at 1.4 VS1, apply 
takeoff power rapidly while maintaining the same airspeed.
    (5) With power off, landing gear and flaps extended, and the 
airplane as nearly as possible in trim at VREF, obtain and 
maintain airspeeds between 1.1 VSO, and either 1.7 
VSO or VFE, whichever is lower without requiring 
the application of two-handed control forces exceeding those specified 
in Sec. 23.143(c).
    (6) With maximum takeoff power, landing gear retracted, flaps in the 
takeoff position, and the airplane as nearly as possible in trim at 
VFE appropriate to the takeoff flap position, retract the 
flaps as rapidly as possible while maintaining constant speed.
    (c) At speeds above VMO/MMO, and up to the 
maximum speed shown under Sec. 23.251, a maneuvering capability of 1.5 
g must be demonstrated to provide a margin to recover from upset or 
inadvertent speed increase.
    (d) It must be possible, with a pilot control force of not more than 
10 pounds, to maintain a speed of not more than VREF during a 
power-off glide with landing gear and wing flaps extended, for any 
weight of the airplane, up to and including the maximum weight.
    (e) By using normal flight and power controls, except as otherwise 
noted in paragraphs (e)(1) and (e)(2) of this section, it must be 
possible to establish a zero rate of descent at an attitude suitable for 
a controlled landing without exceeding the operational and structural 
limitations of the airplane, as follows:
    (1) For single-engine and multiengine airplanes, without the use of 
the primary longitudinal control system.
    (2) For multiengine airplanes--
    (i) Without the use of the primary directional control; and
    (ii) If a single failure of any one connecting or transmitting link 
would affect both the longitudinal and directional primary control 
system, without the primary longitudinal and directional control system.

[Doc. No. 26269, 58 FR 42157, Aug. 6, 1993; Amdt. 23-45, 58 FR 51970, 
Oct. 5, 1993, as amended by Amdt. 23-50, 61 FR 5188, Feb. 9, 1996]



Sec. 23.147  Directional and lateral control.

    (a) For each multiengine airplane, it must be possible, while 
holding the wings level within five degrees, to make sudden changes in 
heading safely in both directions. This ability must be shown at 1.4 
VS1 with heading changes up to 15 degrees, except that the 
heading change at which the rudder force corresponds to the limits 
specified in Sec. 23.143 need not be exceeded, with the--
    (1) Critical engine inoperative and its propeller in the minimum 
drag position;
    (2) Remaining engines at maximum continuous power;
    (3) Landing gear--
    (i) Retracted; and
    (ii) Extended; and
    (4) Flaps retracted.
    (b) For each multiengine airplane, it must be possible to regain 
full control of the airplane without exceeding a bank angle of 45 
degrees, reaching a dangerous attitude or encountering dangerous 
characteristics, in the event of a sudden and complete failure of the 
critical engine, making allowance for a delay of two seconds in the 
initiation of recovery action appropriate to the situation, with the 
airplane initially in trim, in the following condition:
    (1) Maximum continuous power on each engine;
    (2) The wing flaps retracted;
    (3) The landing gear retracted;
    (4) A speed equal to that at which compliance with Sec. 23.69(a) 
has been shown; and

[[Page 199]]

    (5) All propeller controls in the position at which compliance with 
Sec. 23.69(a) has been shown.
    (c) For all airplanes, it must be shown that the airplane is safely 
controllable without the use of the primary lateral control system in 
any all-engine configuration(s) and at any speed or altitude within the 
approved operating envelope. It must also be shown that the airplane's 
flight characteristics are not impaired below a level needed to permit 
continued safe flight and the ability to maintain attitudes suitable for 
a controlled landing without exceeding the operational and structural 
limitations of the airplane. If a single failure of any one connecting 
or transmitting link in the lateral control system would also cause the 
loss of additional control system(s), compliance with the above 
requirement must be shown with those additional systems also assumed to 
be inoperative.

[Doc. No. 27807, 61 FR 5188, Feb. 9, 1996]



Sec. 23.149  Minimum control speed.

    (a) VMC is the calibrated airspeed at which, when the 
critical engine is suddenly made inoperative, it is possible to maintain 
control of the airplane with that engine still inoperative, and 
thereafter maintain straight flight at the same speed with an angle of 
bank of not more than 5 degrees. The method used to simulate critical 
engine failure must represent the most critical mode of powerplant 
failure expected in service with respect to controllability.
    (b) VMC for takeoff must not exceed 1.2 VS1, 
where VS1 is determined at the maximum takeoff weight. 
VMC must be determined with the most unfavorable weight and 
center of gravity position and with the airplane airborne and the ground 
effect negligible, for the takeoff configuration(s) with--
    (1) Maximum available takeoff power initially on each engine;
    (2) The airplane trimmed for takeoff;
    (3) Flaps in the takeoff position(s);
    (4) Landing gear retracted; and
    (5) All propeller controls in the recommended takeoff position 
throughout.
    (c) For all airplanes except reciprocating engine-powered airplanes 
of 6,000 pounds or less maximum weight, the conditions of paragraph (a) 
of this section must also be met for the landing configuration with--
    (1) Maximum available takeoff power initially on each engine;
    (2) The airplane trimmed for an approach, with all engines 
operating, at VREF, at an approach gradient equal to the 
steepest used in the landing distance demonstration of Sec. 23.75;
    (3) Flaps in the landing position;
    (4) Landing gear extended; and
    (5) All propeller controls in the position recommended for approach 
with all engines operating.
    (d) A minimum speed to intentionally render the critical engine 
inoperative must be established and designated as the safe, intentional, 
one-engine-inoperative speed, VSSE.
    (e) At VMC, the rudder pedal force required to maintain 
control must not exceed 150 pounds and it must not be necessary to 
reduce power of the operative engine(s). During the maneuver, the 
airplane must not assume any dangerous attitude and it must be possible 
to prevent a heading change of more than 20 degrees.
    (f) At the option of the applicant, to comply with the requirements 
of Sec. 23.51(c)(1), VMCG may be determined. VMCG 
is the minimum control speed on the ground, and is the calibrated 
airspeed during the takeoff run at which, when the critical engine is 
suddenly made inoperative, it is possible to maintain control of the 
airplane using the rudder control alone (without the use of nosewheel 
steering), as limited by 150 pounds of force, and using the lateral 
control to the extent of keeping the wings level to enable the takeoff 
to be safely continued. In the determination of VMCG, 
assuming that the path of the airplane accelerating with all engines 
operating is along the centerline of the runway, its path from the point 
at which the critical engine is made inoperative to the point at which 
recovery to a direction parallel to the centerline is completed may not 
deviate more than 30 feet laterally from the centerline at any point. 
VMCG must be established with--
    (1) The airplane in each takeoff configuration or, at the option of 
the applicant, in the most critical takeoff configuration;

[[Page 200]]

    (2) Maximum available takeoff power on the operating engines;
    (3) The most unfavorable center of gravity;
    (4) The airplane trimmed for takeoff; and
    (5) The most unfavorable weight in the range of takeoff weights.

[Doc. No. 27807, 61 FR 5189, Feb. 9, 1996]



Sec. 23.151  Acrobatic maneuvers.

    Each acrobatic and utility category airplane must be able to perform 
safely the acrobatic maneuvers for which certification is requested. 
Safe entry speeds for these maneuvers must be determined.



Sec. 23.153  Control during landings.

    It must be possible, while in the landing configuration, to safely 
complete a landing without exceeding the one-hand control force limits 
specified in Sec. 23.143(c) following an approach to land--
    (a) At a speed of VREF minus 5 knots;
    (b) With the airplane in trim, or as nearly as possible in trim and 
without the trimming control being moved throughout the maneuver;
    (c) At an approach gradient equal to the steepest used in the 
landing distance demonstration of Sec. 23.75; and
    (d) With only those power changes, if any, that would be made when 
landing normally from an approach at VREF.

[Doc. No. 27807, 61 FR 5189, Feb. 9, 1996]



Sec. 23.155  Elevator control force in maneuvers.

    (a) The elevator control force needed to achieve the positive limit 
maneuvering load factor may not be less than:
    (1) For wheel controls, W/100 (where W is the maximum weight) or 20 
pounds, whichever is greater, except that it need not be greater than 50 
pounds; or
    (2) For stick controls, W/140 (where W is the maximum weight) or 15 
pounds, whichever is greater, except that it need not be greater than 35 
pounds.
    (b) The requirement of paragraph (a) of this section must be met at 
75 percent of maximum continuous power for reciprocating engines, or the 
maximum continuous power for turbine engines, and with the wing flaps 
and landing gear retracted--
    (1) In a turn, with the trim setting used for wings level flight at 
VO; and
    (2) In a turn with the trim setting used for the maximum wings level 
flight speed, except that the speed may not exceed VNE or 
VMO/MMO, whichever is appropriate.
    (c) There must be no excessive decrease in the gradient of the curve 
of stick force versus maneuvering load factor with increasing load 
factor.

[Amdt. 23-14, 38 FR 31819, Nov. 19, 1973; 38 FR 32784, Nov. 28, 1973, as 
amended by Amdt. 23-45, 58 FR 42158, Aug. 6, 1993; Amdt. 23-50, 61 FR 
5189 Feb. 9, 1996]



Sec. 23.157  Rate of roll.

    (a) Takeoff. It must be possible, using a favorable combination of 
controls, to roll the airplane from a steady 30-degree banked turn 
through an angle of 60 degrees, so as to reverse the direction of the 
turn within:
    (1) For an airplane of 6,000 pounds or less maximum weight, 5 
seconds from initiation of roll; and
    (2) For an airplane of over 6,000 pounds maximum weight,

(W+500)/1,300


seconds, but not more than 10 seconds, where W is the weight in pounds.
    (b) The requirement of paragraph (a) of this section must be met 
when rolling the airplane in each direction with--
    (1) Flaps in the takeoff position;
    (2) Landing gear retracted;
    (3) For a single-engine airplane, at maximum takeoff power; and for 
a multiengine airplane with the critical engine inoperative and the 
propeller in the minimum drag position, and the other engines at maximum 
takeoff power; and
    (4) The airplane trimmed at a speed equal to the greater of 1.2 
VS1 or 1.1 VMC, or as nearly as possible in trim 
for straight flight.
    (c) Approach. It must be possible, using a favorable combination of 
controls, to roll the airplane from a steady 30-degree banked turn 
through an angle of 60 degrees, so as to reverse the direction of the 
turn within:

[[Page 201]]

    (1) For an airplane of 6,000 pounds or less maximum weight, 4 
seconds from initiation of roll; and
    (2) For an airplane of over 6,000 pounds maximum weight,

(W+2,800)/2,200


seconds, but not more than 7 seconds, where W is the weight in pounds.
    (d) The requirement of paragraph (c) of this section must be met 
when rolling the airplane in each direction in the following 
conditions--
    (1) Flaps in the landing position(s);
    (2) Landing gear extended;
    (3) All engines operating at the power for a 3 degree approach; and
    (4) The airplane trimmed at VREF.

[Amdt. 23-14, 38 FR 31819, Nov. 19, 1973, as amended by Amdt. 23-45, 58 
FR 42158, Aug. 6, 1993; Amdt. 23-50, 61 FR 5189, Feb. 9, 1996]

                                  Trim



Sec. 23.161  Trim.

    (a) General. Each airplane must meet the trim requirements of this 
section after being trimmed and without further pressure upon, or 
movement of, the primary controls or their corresponding trim controls 
by the pilot or the automatic pilot. In addition, it must be possible, 
in other conditions of loading, configuration, speed and power to ensure 
that the pilot will not be unduly fatigued or distracted by the need to 
apply residual control forces exceeding those for prolonged application 
of Sec. 23.143(c). This applies in normal operation of the airplane 
and, if applicable, to those conditions associated with the failure of 
one engine for which performance characteristics are established.
    (b) Lateral and directional trim. The airplane must maintain lateral 
and directional trim in level flight with the landing gear and wing 
flaps retracted as follows:
    (1) For normal, utility, and acrobatic category airplanes, at a 
speed of 0.9 VH, VC, or VMO/
MO, whichever is lowest; and
    (2) For commuter category airplanes, at all speeds from 1.4 
VS1 to the lesser of VH or VMO/
MMO.
    (c) Longitudinal trim. The airplane must maintain longitudinal trim 
under each of the following conditions:
    (1) A climb with--
    (i) Takeoff power, landing gear retracted, wing flaps in the takeoff 
position(s), at the speeds used in determining the climb performance 
required by Sec. 23.65; and
    (ii) Maximum continuous power at the speeds and in the configuration 
used in determining the climb performance required by Sec. 23.69(a).
    (2) Level flight at all speeds from the lesser of VH and 
either VNO or VMO/MMO (as appropriate), 
to 1.4 VS1, with the landing gear and flaps retracted.
    (3) A descent at VNO or VMO/MMO, 
whichever is applicable, with power off and with the landing gear and 
flaps retracted.
    (4) Approach with landing gear extended and with--
    (i) A 3 degree angle of descent, with flaps retracted and at a speed 
of 1.4 VS1;
    (ii) A 3 degree angle of descent, flaps in the landing position(s) 
at VREF; and
    (iii) An approach gradient equal to the steepest used in the landing 
distance demonstrations of Sec. 23.75, flaps in the landing position(s) 
at VREF.
    (d) In addition, each multiple airplane must maintain longitudinal 
and directional trim, and the lateral control force must not exceed 5 
pounds at the speed used in complying with Sec. 23.67(a), (b)(2), or 
(c)(3), as appropriate, with--
    (1) The critical engine inoperative, and if applicable, its 
propeller in the minimum drag position;
    (2) The remaining engines at maximum continuous power;
    (3) The landing gear retracted;
    (4) Wing flaps retracted; and
    (5) An angle of bank of not more than five degrees.
    (e) In addition, each commuter category airplane for which, in the 
determination of the takeoff path in accordance with Sec. 23.57, the 
climb in the takeoff configuration at V2 extends beyond 400 
feet above the takeoff surface, it must be possible to reduce the 
longitudinal and lateral control forces to 10 pounds and 5 pounds, 
respectively, and the directional control force must not exceed 50 
pounds at V2 with--
    (1) The critical engine inoperative and its propeller in the minimum 
drag position;
    (2) The remaining engine(s) at takeoff power;

[[Page 202]]

    (3) Landing gear retracted;
    (4) Wing flaps in the takeoff position(s); and
    (5) An angle of bank not exceeding 5 degrees.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-21, 
43 FR 2318, Jan. 16, 1978; Amdt. 23-34, 52 FR 1828, Jan. 15, 1987; Amdt. 
23-42, 56 FR 351, Jan. 3, 1991; 56 FR 5455, Feb. 11, 1991; Amdt. 23-50, 
61 FR 5189, Feb. 9, 1996]

                                Stability



Sec. 23.171  General.

    The airplane must be longitudinally, directionally, and laterally 
stable under Sec. Sec. 23.173 through 23.181. In addition, the airplane 
must show suitable stability and control ``feel'' (static stability) in 
any condition normally encountered in service, if flight tests show it 
is necessary for safe operation.



Sec. 23.173  Static longitudinal stability.

    Under the conditions specified in Sec. 23.175 and with the airplane 
trimmed as indicated, the characteristics of the elevator control forces 
and the friction within the control system must be as follows:
    (a) A pull must be required to obtain and maintain speeds below the 
specified trim speed and a push required to obtain and maintain speeds 
above the specified trim speed. This must be shown at any speed that can 
be obtained, except that speeds requiring a control force in excess of 
40 pounds or speeds above the maximum allowable speed or below the 
minimum speed for steady unstalled flight, need not be considered.
    (b) The airspeed must return to within the tolerances specified for 
applicable categories of airplanes when the control force is slowly 
released at any speed within the speed range specified in paragraph (a) 
of this section. The applicable tolerances are--
    (1) The airspeed must return to within plus or minus 10 percent of 
the original trim airspeed; and
    (2) For commuter category airplanes, the airspeed must return to 
within plus or minus 7.5 percent of the original trim airspeed for the 
cruising condition specified in Sec. 23.175(b).
    (c) The stick force must vary with speed so that any substantial 
speed change results in a stick force clearly perceptible to the pilot.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31820 Nov. 19, 1973; Amdt. 23-34, 52 FR 1828, Jan. 15, 1987]



Sec. 23.175  Demonstration of static longitudinal stability.

    Static longitudinal stability must be shown as follows:
    (a) Climb. The stick force curve must have a stable slope at speeds 
between 85 and 115 percent of the trim speed, with--
    (1) Flaps retracted;
    (2) Landing gear retracted;
    (3) Maximum continuous power; and
    (4) The airplane trimmed at the speed used in determining the climb 
performance required by Sec. 23.69(a).
    (b) Cruise. With flaps and landing gear retracted and the airplane 
in trim with power for level flight at representative cruising speeds at 
high and low altitudes, including speeds up to VNO or 
VMO/MMO, as appropriate, except that the speed 
need not exceed VH--
    (1) For normal, utility, and acrobatic category airplanes, the stick 
force curve must have a stable slope at all speeds within a range that 
is the greater of 15 percent of the trim speed plus the resulting free 
return speed range, or 40 knots plus the resulting free return speed 
range, above and below the trim speed, except that the slope need not be 
stable--
    (i) At speeds less than 1.3 VS1; or
    (ii) For airplanes with VNE established under Sec. 
23.1505(a), at speeds greater than VNE; or
    (iii) For airplanes with VMO/MMO established 
under Sec. 23.1505(c), at speeds greater than VFC/
MFC.
    (2) For commuter category airplanes, the stick force curve must have 
a stable slope at all speeds within a range of 50 knots plus the 
resulting free return speed range, above and below the trim speed, 
except that the slope need not be stable--
    (i) At speeds less than 1.4 VS1; or
    (ii) At speeds greater than VFC/MFC; or
    (iii) At speeds that require a stick force greater than 50 pounds.

[[Page 203]]

    (c) Landing. The stick force curve must have a stable slope at 
speeds between 1.1 VS1 and 1.8 VS1 with--
    (1) Flaps in the landing position;
    (2) Landing gear extended; and
    (3) The airplane trimmed at--
    (i) VREF, or the minimum trim speed if higher, with power 
off; and
    (ii) VREF with enough power to maintain a 3 degree angle 
of descent.

[Doc. No. 27807, 61 FR 5190, Feb. 9, 1996]



Sec. 23.177  Static directional and lateral stability.

    (a)(1) The static directional stability, as shown by the tendency to 
recover from a wings level sideslip with the rudder free, must be 
positive for any landing gear and flap position appropriate to the 
takeoff, climb, cruise, approach, and landing configurations. This must 
be shown with symmetrical power up to maximum continuous power, and at 
speeds from 1.2 VS1 up to VFE, VLE, 
VNO, VFC/MFC, whichever is appropriate.
    (2) The angle of sideslip for these tests must be appropriate to the 
type of airplane. The rudder pedal force must not reverse at larger 
angles of sideslip, up to that at which full rudder is used or a control 
force limit in Sec. 23.143 is reached, whichever occurs first, and at 
speeds from 1.2 VS1 to VO.
    (b)(1) The static lateral stability, as shown by the tendency to 
raise the low wing in a sideslip with the aileron controls free, may not 
be negative for any landing gear and flap position appropriate to the 
takeoff, climb, cruise, approach, and landing configurations. This must 
be shown with symmetrical power from idle up to 75 percent of maximum 
continuous power at speeds from 1.2 VS1 in the takeoff 
configuration(s) and at speeds from 1.3 VS1 in other 
configurations, up to the maximum allowable airspeed for the 
configuration being investigated (VFE, VLE, 
VNO, VFC/MFC, whichever is appropriate) 
in the takeoff, climb, cruise, descent, and approach configurations. For 
the landing configuration, the power must be that necessary to maintain 
a 3-degree angle of descent in coordinated flight.
    (2) The static lateral stability may not be negative at 1.2 
VS1 in the takeoff configuration, or at 1.3 VS1 in 
other configurations.
    (3) The angel of sideslip for these tests must be appropriate to the 
type of airplane, but in no case may the constant heading sideslip angle 
be less than that obtainable with a 10 degree bank or, if less, the 
maximum bank angle obtainable with full rudder deflection or 150 pound 
rudder force.
    (c) Paragraph (b) of this section does not apply to acrobatic 
category airplanes certificated for inverted flight.
    (d)(1) In straight, steady slips at 1.2 VS1 for any 
landing gear and flap position appropriate to the takeoff, climb, 
cruise, approach, and landing configurations, and for any symmetrical 
power conditions up to 50 percent of maximum continuous power, the 
aileron and rudder control movements and forces must increase steadily, 
but not necessarily in constant proportion, as the angle of sideslip is 
increased up to the maximum appropriate to the type of airplane.
    (2) At larger slip angles, up to the angle at which the full rudder 
or aileron control is used or a control force limit contained in Sec. 
23.143 is reached, the aileron and rudder control movements and forces 
may not reverse as the angle of sideslip is increased.
    (3) Rapid entry into, and recovery from, a maximum sideslip 
considered appropriate for the airplane may not result in uncontrollable 
flight characteristics.

[Doc. No. 27807, 61 FR 5190, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75754, Dec. 2, 2011]



Sec. 23.181  Dynamic stability.

    (a) Any short period oscillation not including combined lateral-
directional oscillations occurring between the stalling speed and the 
maximum allowable speed appropriate to the configuration of the airplane 
must be heavily damped with the primary controls--
    (1) Free; and
    (2) In a fixed position.
    (b) Any combined lateral-directional oscillations (Dutch roll) 
occurring between the stalling speed and the maximum allowable speed 
(VFE, VLE, VN0, VFC/
MFC) appropriate to the configuration of the airplane with 
the primary

[[Page 204]]

controls in both free and fixed position, must be damped to \1/10\ 
amplitude in:
    (1) Seven (7) cycles below 18,000 feet and
    (2) Thirteen (13) cycles from 18,000 feet to the certified maximum 
altitude.
    (c) If it is determined that the function of a stability 
augmentation system, reference Sec. 23.672, is needed to meet the 
flight characteristic requirements of this part, the primary control 
requirements of paragraphs (a)(2) and (b)(2) of this section are not 
applicable to the tests needed to verify the acceptability of that 
system.
    (d) During the conditions as specified in Sec. 23.175, when the 
longitudinal control force required to maintain speeds differing from 
the trim speed by at least plus and minus 15 percent is suddenly 
released, the response of the airplane must not exhibit any dangerous 
characteristics nor be excessive in relation to the magnitude of the 
control force released. Any long-period oscillation of flight path, 
phugoid oscillation, that results must not be so unstable as to increase 
the pilot's workload or otherwise endanger the airplane.

[Amdt. 23-21, 43 FR 2318, Jan. 16, 1978, as amended by Amdt. 23-45, 58 
FR 42158, Aug. 6, 1993; Amdt. 23-62, 76 FR 75755, Dec. 2, 2011]

                                 Stalls



Sec. 23.201  Wings level stall.

    (a) It must be possible to produce and to correct roll by unreversed 
use of the rolling control and to produce and to correct yaw by 
unreversed use of the directional control, up to the time the airplane 
stalls.
    (b) The wings level stall characteristics must be demonstrated in 
flight as follows. Starting from a speed at least 10 knots above the 
stall speed, the elevator control must be pulled back so that the rate 
of speed reduction will not exceed one knot per second until a stall is 
produced, as shown by either:
    (1) An uncontrollable downward pitching motion of the airplane;
    (2) A downward pitching motion of the airplane that results from the 
activation of a stall avoidance device (for example, stick pusher); or
    (3) The control reaching the stop.
    (c) Normal use of elevator control for recovery is allowed after the 
downward pitching motion of paragraphs (b)(1) or (b)(2) of this section 
has unmistakably been produced, or after the control has been held 
against the stop for not less than the longer of two seconds or the time 
employed in the minimum steady slight speed determination of Sec. 
23.49.
    (d) During the entry into and the recovery from the maneuver, it 
must be possible to prevent more than 15 degrees of roll or yaw by the 
normal use of controls except as provided for in paragraph (e) of this 
section.
    (e) For airplanes approved with a maximum operating altitude at or 
above 25,000 feet during the entry into and the recovery from stalls 
performed at or above 25,000 feet, it must be possible to prevent more 
than 25 degrees of roll or yaw by the normal use of controls.
    (f) Compliance with the requirements of this section must be shown 
under the following conditions:
    (1) Wing flaps: Retracted, fully extended, and each intermediate 
normal operating position, as appropriate for the phase of flight.
    (2) Landing gear: Retracted and extended as appropriate for the 
altitude.
    (3) Cowl flaps: Appropriate to configuration.
    (4) Spoilers/speedbrakes: Retracted and extended unless they have no 
measureable effect at low speeds.
    (5) Power:
    (i) Power/Thrust off; and
    (ii) For reciprocating engine powered airplanes: 75 percent of 
maximum continuous power. However, if the power-to-weight ratio at 75 
percent of maximum continuous power results in nose-high attitudes 
exceeding 30 degrees, the test may be carried out with the power 
required for level flight in the landing configuration at maximum 
landing weight and a speed of 1.4 VSO, except that the power 
may not be less than 50 percent of maximum continuous power; or
    (iii) For turbine engine powered airplanes: The maximum engine 
thrust, except that it need not exceed the thrust necessary to maintain 
level flight at 1.5 VS1 (where VS1 corresponds to 
the stalling speed with flaps in the approach position, the landing gear 
retracted, and maximum landing weight).

[[Page 205]]

    (6) Trim: At 1.5 VS1 or the minimum trim speed, whichever 
is higher.
    (7) Propeller: Full increase r.p.m. position for the power off 
condition.

[Doc. No. 27807, 61 FR 5191, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75755, Dec. 2, 2011]



Sec. 23.203  Turning flight and accelerated turning stalls.

    Turning flight and accelerated turning stalls must be demonstrated 
in tests as follows:
    (a) Establish and maintain a coordinated turn in a 30 degree bank. 
Reduce speed by steadily and progressively tightening the turn with the 
elevator until the airplane is stalled, as defined in Sec. 23.201(b). 
The rate of speed reduction must be constant, and--
    (1) For a turning flight stall, may not exceed one knot per second; 
and
    (2) For an accelerated turning stall, be 3 to 5 knots per second 
with steadily increasing normal acceleration.
    (b) After the airplane has stalled, as defined in Sec. 23.201(b), 
it must be possible to regain wings level flight by normal use of the 
flight controls, but without increasing power and without--
    (1) Excessive loss of altitude;
    (2) Undue pitchup;
    (3) Uncontrollable tendency to spin;
    (4) Exceeding a bank angle of 60 degrees in the original direction 
of the turn or 30 degrees in the opposite direction in the case of 
turning flight stalls;
    (5) Exceeding a bank angle of 90 degrees in the original direction 
of the turn or 60 degrees in the opposite direction in the case of 
accelerated turning stalls; and
    (6) Exceeding the maximum permissible speed or allowable limit load 
factor.
    (c) Compliance with the requirements of this section must be shown 
under the following conditions:
    (1) Wings flaps: Retracted, fully extended, and each intermediate 
normal operating position as appropriate for the phase of flight.
    (2) Landing gear: Retracted and extended as appropriate for the 
altitude.
    (3) Cowl flaps: Appropriate to configuration.
    (4) Spoilers/speedbrakes: Retracted and extended unless they have no 
measureable effect at low speeds.
    (5) Power:
    (i) Power/Thrust off; and
    (ii) For reciprocating engine powered airplanes: 75 percent of 
maximum continuous power. However, if the power-to-weight ratio at 75 
percent of maximum continuous power results in nose-high attitudes 
exceeding 30 degrees, the test may be carried out with the power 
required for level flight in the landing configuration at maximum 
landing weight and a speed of 1.4 VSO, except that the power 
may not be less than 50 percent of maximum continuous power; or
    (iii) For turbine engine powered airplanes: The maximum engine 
thrust, except that it need not exceed the thrust necessary to maintain 
level flight at 1.5 VS1 (where VS1 corresponds to 
the stalling speed with flaps in the approach position, the landing gear 
retracted, and maximum landing weight).
    (6) Trim: The airplane trimmed at 1.5 VS1.
    (7) Propeller: Full increase rpm position for the power off 
condition.

[Amdt. 23-14, 38 FR 31820, Nov. 19, 1973, as amended by Amdt. 23-45, 58 
FR 42159, Aug. 6, 1993; Amdt. 23-50, 61 FR 5191, Feb. 9, 1996; Amdt. 23-
62, 76 FR 75755, Dec. 2, 2011]



Sec. 23.207  Stall warning.

    (a) There must be a clear and distinctive stall warning, with the 
flaps and landing gear in any normal position, in straight and turning 
flight.
    (b) The stall warning may be furnished either through the inherent 
aerodynamic qualities of the airplane or by a device that will give 
clearly distinguishable indications under expected conditions of flight. 
However, a visual stall warning device that requires the attention of 
the crew within the cockpit is not acceptable by itself.
    (c) During the stall tests required by Sec. 23.201(b) and Sec. 
23.203(a)(1), the stall warning must begin at a speed exceeding the 
stalling speed by a margin of not less than 5 knots and must continue 
until the stall occurs.
    (d) When following procedures furnished in accordance with Sec. 
23.1585, the stall warning must not occur during a takeoff with all 
engines operating, a takeoff continued with one engine inoperative, or 
during an approach to landing.

[[Page 206]]

    (e) During the stall tests required by Sec. 23.203(a)(2), the stall 
warning must begin sufficiently in advance of the stall for the stall to 
be averted by pilot action taken after the stall warning first occurs.
    (f) For acrobatic category airplanes, an artificial stall warning 
may be mutable, provided that it is armed automatically during takeoff 
and rearmed automatically in the approach configuration.

[Amdt. 23-7, 34 FR 13087, Aug. 13, 1969, as amended by Amdt. 23-45, 58 
FR 42159, Aug. 6, 1993; Amdt. 23-50, 61 FR 5191, Feb. 9, 1996]

                                Spinning



Sec. 23.221  Spinning.

    (a) Normal category airplanes. A single-engine, normal category 
airplane must be able to recover from a one-turn spin or a three-second 
spin, whichever takes longer, in not more than one additional turn after 
initiation of the first control action for recovery, or demonstrate 
compliance with the optional spin resistant requirements of this 
section.
    (1) The following apply to one turn or three second spins:
    (i) For both the flaps-retracted and flaps-extended conditions, the 
applicable airspeed limit and positive limit maneuvering load factor 
must not be exceeded;
    (ii) No control forces or characteristic encountered during the spin 
or recovery may adversely affect prompt recovery;
    (iii) It must be impossible to obtain unrecoverable spins with any 
use of the flight or engine power controls either at the entry into or 
during the spin; and
    (iv) For the flaps-extended condition, the flaps may be retracted 
during the recovery but not before rotation has ceased.
    (2) At the applicant's option, the airplane may be demonstrated to 
be spin resistant by the following:
    (i) During the stall maneuver contained in Sec. 23.201, the pitch 
control must be pulled back and held against the stop. Then, using 
ailerons and rudders in the proper direction, it must be possible to 
maintain wings-level flight within 15 degrees of bank and to roll the 
airplane from a 30 degree bank in one direction to a 30 degree bank in 
the other direction;
    (ii) Reduce the airplane speed using pitch control at a rate of 
approximately one knot per second until the pitch control reaches the 
stop; then, with the pitch control pulled back and held against the 
stop, apply full rudder control in a manner to promote spin entry for a 
period of seven seconds or through a 360 degree heading change, 
whichever occurs first. If the 360 degree heading change is reached 
first, it must have taken no fewer than four seconds. This maneuver must 
be performed first with the ailerons in the neutral position, and then 
with the ailerons deflected opposite the direction of turn in the most 
adverse manner. Power and airplane configuration must be set in 
accordance with Sec. 23.201(e) without change during the maneuver. At 
the end of seven seconds or a 360 degree heading change, the airplane 
must respond immediately and normally to primary flight controls applied 
to regain coordinated, unstalled flight without reversal of control 
effect and without exceeding the temporary control forces specified by 
Sec. 23.143(c); and
    (iii) Compliance with Sec. Sec. 23.201 and 23.203 must be 
demonstrated with the airplane in uncoordinated flight, corresponding to 
one ball width displacement on a slip-skid indicator, unless one ball 
width displacement cannot be obtained with full rudder, in which case 
the demonstration must be with full rudder applied.
    (b) Utility category airplanes. A utility category airplane must 
meet the requirements of paragraph (a) of this section. In addition, the 
requirements of paragraph (c) of this section and Sec. 23.807(b)(7) 
must be met if approval for spinning is requested.
    (c) Acrobatic category airplanes. An acrobatic category airplane 
must meet the spin requirements of paragraph (a) of this section and 
Sec. 23.807(b)(6). In addition, the following requirements must be met 
in each configuration for which approval for spinning is requested:
    (1) The airplane must recover from any point in a spin up to and 
including six turns, or any greater number of turns for which 
certification is requested, in not more than one and one-

[[Page 207]]

half additional turns after initiation of the first control action for 
recovery. However, beyond three turns, the spin may be discontinued if 
spiral characteristics appear.
    (2) The applicable airspeed limits and limit maneuvering load 
factors must not be exceeded. For flaps-extended configurations for 
which approval is requested, the flaps must not be retracted during the 
recovery.
    (3) It must be impossible to obtain unrecoverable spins with any use 
of the flight or engine power controls either at the entry into or 
during the spin.
    (4) There must be no characteristics during the spin (such as 
excessive rates of rotation or extreme oscillatory motion) that might 
prevent a successful recovery due to disorientation or incapacitation of 
the pilot.

[Doc. No. 27807, 61 FR 5191, Feb. 9, 1996]

                Ground and Water Handling Characteristics



Sec. 23.231  Longitudinal stability and control.

    (a) A landplane may have no uncontrollable tendency to nose over in 
any reasonably expected operating condition, including rebound during 
landing or takeoff. Wheel brakes must operate smoothly and may not 
induce any undue tendency to nose over.
    (b) A seaplane or amphibian may not have dangerous or uncontrollable 
porpoising characteristics at any normal operating speed on the water.



Sec. 23.233  Directional stability and control.

    (a) A 90 degree cross-component of wind velocity, demonstrated to be 
safe for taxiing, takeoff, and landing must be established and must be 
not less than 0.2 VSO.
    (b) The airplane must be satisfactorily controllable in power-off 
landings at normal landing speed, without using brakes or engine power 
to maintain a straight path until the speed has decreased to at least 50 
percent of the speed at touchdown.
    (c) The airplane must have adequate directional control during 
taxiing.
    (d) Seaplanes must demonstrate satisfactory directional stability 
and control for water operations up to the maximum wind velocity 
specified in paragraph (a) of this section.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42159, Aug. 6, 1993; Amdt. 23-50, 61 FR 5192, Feb. 9, 1996]



Sec. 23.235  Operation on unpaved surfaces.

    The airplane must be demonstrated to have satisfactory 
characteristics and the shock-absorbing mechanism must not damage the 
structure of the airplane when the airplane is taxied on the roughest 
ground that may reasonably be expected in normal operation and when 
takeoffs and landings are performed on unpaved runways having the 
roughest surface that may reasonably be expected in normal operation.

[Doc. No. 27807, 61 FR 5192, Feb. 9, 1996]



Sec. 23.237  Operation on water.

    A wave height, demonstrated to be safe for operation, and any 
necessary water handling procedures for seaplanes and amphibians must be 
established.

[Doc. No. 27807, 61 FR 5192, Feb. 9, 1996]



Sec. 23.239  Spray characteristics.

    Spray may not dangerously obscure the vision of the pilots or damage 
the propellers or other parts of a seaplane or amphibian at any time 
during taxiing, takeoff, and landing.

                    Miscellaneous Flight Requirements



Sec. 23.251  Vibration and buffeting.

    (a) There must be no vibration or buffeting severe enough to result 
in structural damage, and each part of the airplane must be free from 
excessive vibration, under any appropriate speed and power conditions up 
to VD/MD, or VDF/MDF for 
turbojets. In addition, there must be no buffeting in any normal flight 
condition, including configuration changes during cruise, severe enough 
to interfere with the satisfactory control of the airplane or cause 
excessive fatigue to the flight crew. Stall warning buffeting within 
these limits is allowable.
    (b) There must be no perceptible buffeting condition in the cruise 
configuration in straight flight at any speed

[[Page 208]]

up to VMO/MMO, except stall buffeting, which is 
allowable.
    (c) For airplanes with MD greater than M 0.6 or a maximum 
operating altitude greater than 25,000 feet, the positive maneuvering 
load factors at which the onset of perceptible buffeting occurs must be 
determined with the airplane in the cruise configuration for the ranges 
of airspeed or Mach number, weight, and altitude for which the airplane 
is to be certificated. The envelopes of load factor, speed, altitude, 
and weight must provide a sufficient range of speeds and load factors 
for normal operations. Probable inadvertent excursions beyond the 
boundaries of the buffet onset envelopes may not result in unsafe 
conditions.

[Amdt. 23-62, 76 FR 75755, Dec. 2, 2011]



Sec. 23.253  High speed characteristics.

    If a maximum operating speed VMO/MMO is 
established under Sec. 23.1505(c), the following speed increase and 
recovery characteristics must be met:
    (a) Operating conditions and characteristics likely to cause 
inadvertent speed increases (including upsets in pitch and roll) must be 
simulated with the airplane trimmed at any likely speed up to 
VMO/MMO. These conditions and characteristics 
include gust upsets, inadvertent control movements, low stick force 
gradients in relation to control friction, passenger movement, leveling 
off from climb, and descent from Mach to airspeed limit altitude.
    (b) Allowing for pilot reaction time after occurrence of the 
effective inherent or artificial speed warning specified in Sec. 
23.1303, it must be shown that the airplane can be recovered to a normal 
attitude and its speed reduced to VMO/MMO, 
without--
    (1) Exceptional piloting strength or skill;
    (2) Exceeding VD/MD, or VDF/
MDF for turbojets, the maximum speed shown under Sec. 
23.251, or the structural limitations; and
    (3) Buffeting that would impair the pilot's ability to read the 
instruments or to control the airplane for recovery.
    (c) There may be no control reversal about any axis at any speed up 
to the maximum speed shown under Sec. 23.251. Any reversal of elevator 
control force or tendency of the airplane to pitch, roll, or yaw must be 
mild and readily controllable, using normal piloting techniques.
    (d) Maximum speed for stability characteristics, VFC/MFC. 
VFC/MFC may not be less than a speed midway 
between VMO/MMO and VDF/MDF 
except that, for altitudes where Mach number is the limiting factor, 
MFC need not exceed the Mach number at which effective speed 
warning occurs.

[Amdt. 23-7, 34 FR 13087, Aug. 13, 1969; as amended by Amdt. 23-26, 45 
FR 60170, Sept. 11, 1980; Amdt. 23-45, 58 FR 42160, Aug. 6, 1993; Amdt. 
23-50, 61 FR 5192, Feb. 9, 1996; Amdt. 23-62, 76 FR 75755, Dec. 2, 2011]



Sec. 23.255  Out of trim characteristics.

    For airplanes with an MD greater than M 0.6 and that 
incorporate a trimmable horizontal stabilizer, the following 
requirements for out-of-trim characteristics apply:
    (a) From an initial condition with the airplane trimmed at cruise 
speeds up to VMO/MMO, the airplane must have 
satisfactory maneuvering stability and controllability with the degree 
of out-of-trim in both the airplane nose-up and nose-down directions, 
which results from the greater of the following:
    (1) A three-second movement of the longitudinal trim system at its 
normal rate for the particular flight condition with no aerodynamic load 
(or an equivalent degree of trim for airplanes that do not have a power-
operated trim system), except as limited by stops in the trim system, 
including those required by Sec. 23.655(b) for adjustable stabilizers; 
or
    (2) The maximum mistrim that can be sustained by the autopilot while 
maintaining level flight in the high speed cruising condition.
    (b) In the out-of-trim condition specified in paragraph (a) of this 
section, when the normal acceleration is varied from +l g to the 
positive and negative values specified in paragraph (c) of this section, 
the following apply:
    (1) The stick force versus g curve must have a positive slope at any 
speed up to and including VFC/MFC; and
    (2) At speeds between VFC/MFC and 
VDF/MDF, the direction of the primary 
longitudinal control force may not reverse.

[[Page 209]]

    (c) Except as provided in paragraphs (d) and (e) of this section, 
compliance with the provisions of paragraph (a) of this section must be 
demonstrated in flight over the acceleration range as follows:
    (1) -1 g to +2.5 g; or
    (2) 0 g to 2.0 g, and extrapolating by an acceptable method to -1 g 
and +2.5 g.
    (d) If the procedure set forth in paragraph (c)(2) of this section 
is used to demonstrate compliance and marginal conditions exist during 
flight test with regard to reversal of primary longitudinal control 
force, flight tests must be accomplished from the normal acceleration at 
which a marginal condition is found to exist to the applicable limit 
specified in paragraph (b)(1) of this section.
    (e) During flight tests required by paragraph (a) of this section, 
the limit maneuvering load factors, prescribed in Sec. Sec. 23.333(b) 
and 23.337, need not be exceeded. In addition, the entry speeds for 
flight test demonstrations at normal acceleration values less than 1 g 
must be limited to the extent necessary to accomplish a recovery without 
exceeding VDF/MDF.
    (f) In the out-of-trim condition specified in paragraph (a) of this 
section, it must be possible from an overspeed condition at 
VDF/MDF to produce at least 1.5 g for recovery by 
applying not more than 125 pounds of longitudinal control force using 
either the primary longitudinal control alone or the primary 
longitudinal control and the longitudinal trim system. If the 
longitudinal trim is used to assist in producing the required load 
factor, it must be shown at VDF/MDF that the 
longitudinal trim can be actuated in the airplane nose-up direction with 
the primary surface loaded to correspond to the least of the following 
airplane nose-up control forces:
    (1) The maximum control forces expected in service, as specified in 
Sec. Sec. 23.301 and 23.397.
    (2) The control force required to produce 1.5 g.
    (3) The control force corresponding to buffeting or other phenomena 
of such intensity that it is a strong deterrent to further application 
of primary longitudinal control force.

[Doc. No. FAA-2009-0738, 76 FR 75755, Dec. 2, 2011]



                           Subpart C_Structure

                                 General



Sec. 23.301  Loads.

    (a) Strength requirements are specified in terms of limit loads (the 
maximum loads to be expected in service) and ultimate loads (limit loads 
multiplied by prescribed factors of safety). Unless otherwise provided, 
prescribed loads are limit loads.
    (b) Unless otherwise provided, the air, ground, and water loads must 
be placed in equilibrium with inertia forces, considering each item of 
mass in the airplane. These loads must be distributed to conservatively 
approximate or closely represent actual conditions. Methods used to 
determine load intensities and distribution on canard and tandem wing 
configurations must be validated by flight test measurement unless the 
methods used for determining those loading conditions are shown to be 
reliable or conservative on the configuration under consideration.
    (c) If deflections under load would significantly change the 
distribution of external or internal loads, this redistribution must be 
taken into account.
    (d) Simplified structural design criteria may be used if they result 
in design loads not less than those prescribed in Sec. Sec. 23.331 
through 23.521. For airplane configurations described in appendix A, 
Sec. 23.1, the design criteria of appendix A of this part are an 
approved equivalent of Sec. Sec. 23.321 through 23.459. If appendix A 
of this part is used, the entire appendix must be substituted for the 
corresponding sections of this part.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. 23-42, 56 FR 
352, Jan. 3, 1991; Amdt. 23-48, 61 FR 5143, Feb. 9, 1996]



Sec. 23.302  Canard or tandem wing configurations.

    The forward structure of a canard or tandem wing configuration must:

[[Page 210]]

    (a) Meet all requirements of subpart C and subpart D of this part 
applicable to a wing; and
    (b) Meet all requirements applicable to the function performed by 
these surfaces.

[Amdt. 23-42, 56 FR 352, Jan. 3, 1991]



Sec. 23.303  Factor of safety.

    Unless otherwise provided, a factor of safety of 1.5 must be used.



Sec. 23.305  Strength and deformation.

    (a) The structure must be able to support limit loads without 
detrimental, permanent deformation. At any load up to limit loads, the 
deformation may not interfere with safe operation.
    (b) The structure must be able to support ultimate loads without 
failure for at least three seconds, except local failures or structural 
instabilities between limit and ultimate load are acceptable only if the 
structure can sustain the required ultimate load for at least three 
seconds. However when proof of strength is shown by dynamic tests 
simulating actual load conditions, the three second limit does not 
apply.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42160, Aug. 6, 1993]



Sec. 23.307  Proof of structure.

    (a) Compliance with the strength and deformation requirements of 
Sec. 23.305 must be shown for each critical load condition. Structural 
analysis may be used only if the structure conforms to those for which 
experience has shown this method to be reliable. In other cases, 
substantiating load tests must be made. Dynamic tests, including 
structural flight tests, are acceptable if the design load conditions 
have been simulated.
    (b) Certain parts of the structure must be tested as specified in 
Subpart D of this part.

                              Flight Loads



Sec. 23.321  General.

    (a) Flight load factors represent the ratio of the aerodynamic force 
component (acting normal to the assumed longitudinal axis of the 
airplane) to the weight of the airplane. A positive flight load factor 
is one in which the aerodynamic force acts upward, with respect to the 
airplane.
    (b) Compliance with the flight load requirements of this subpart 
must be shown--
    (1) At each critical altitude within the range in which the airplane 
may be expected to operate;
    (2) At each weight from the design minimum weight to the design 
maximum weight; and
    (3) For each required altitude and weight, for any practicable 
distribution of disposable load within the operating limitations 
specified in Sec. Sec. 23.1583 through 23.1589.
    (c) When significant, the effects of compressibility must be taken 
into account.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42160, Aug. 6, 1993]



Sec. 23.331  Symmetrical flight conditions.

    (a) The appropriate balancing horizontal tail load must be accounted 
for in a rational or conservative manner when determining the wing loads 
and linear inertia loads corresponding to any of the symmetrical flight 
conditions specified in Sec. Sec. 23.333 through 23.341.
    (b) The incremental horizontal tail loads due to maneuvering and 
gusts must be reacted by the angular inertia of the airplane in a 
rational or conservative manner.
    (c) Mutual influence of the aerodynamic surfaces must be taken into 
account when determining flight loads.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-42, 56 FR 352, Jan. 3, 1991]



Sec. 23.333  Flight envelope.

    (a) General. Compliance with the strength requirements of this 
subpart must be shown at any combination of airspeed and load factor on 
and within the boundaries of a flight envelope (similar to the one in 
paragraph (d) of this section) that represents the envelope of the 
flight loading conditions specified by the maneuvering and gust criteria 
of paragraphs (b) and (c) of this section respectively.

[[Page 211]]

    (b) Maneuvering envelope. Except where limited by maximum (static) 
lift coefficients, the airplane is assumed to be subjected to 
symmetrical maneuvers resulting in the following limit load factors:
    (1) The positive maneuvering load factor specified in Sec. 23.337 
at speeds up to VD;
    (2) The negative maneuvering load factor specified in Sec. 23.337 
at VC; and
    (3) Factors varying linearly with speed from the specified value at 
VC to 0.0 at VD for the normal and commuter category, and -1.0 at VD for 
the acrobatic and utility categories.
    (c) Gust envelope. (1) The airplane is assumed to be subjected to 
symmetrical vertical gusts in level flight. The resulting limit load 
factors must correspond to the conditions determined as follows:
    (i) Positive (up) and negative (down) gusts of 50 f.p.s. at VC must 
be considered at altitudes between sea level and 20,000 feet. The gust 
velocity may be reduced linearly from 50 f.p.s. at 20,000 feet to 25 
f.p.s. at 50,000 feet.
    (ii) Positive and negative gusts of 25 f.p.s. at VD must be 
considered at altitudes between sea level and 20,000 feet. The gust 
velocity may be reduced linearly from 25 f.p.s. at 20,000 feet to 12.5 
f.p.s. at 50,000 feet.
    (iii) In addition, for commuter category airplanes, positive (up) 
and negative (down) rough air gusts of 66 f.p.s. at V[Beta] must be 
considered at altitudes between sea level and 20,000 feet. The gust 
velocity may be reduced linearly from 66 f.p.s. at 20,000 feet to 38 
f.p.s. at 50,000 feet.
    (2) The following assumptions must be made:
    (i) The shape of the gust is--
    [GRAPHIC] [TIFF OMITTED] TC28SE91.000
    
Where--

s=Distance penetrated into gust (ft.);
C=Mean geometric chord of wing (ft.); and
Ude=Derived gust velocity referred to in subparagraph (1) of this 
          section.

    (ii) Gust load factors vary linearly with speed between VC and VD .
    (d) Flight envelope.
    [GRAPHIC] [TIFF OMITTED] TC28SE91.001
    

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13087, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987]

[[Page 212]]



Sec. 23.335  Design airspeeds.

    Except as provided in paragraph (a)(4) of this section, the selected 
design airspeeds are equivalent airspeeds (EAS).
    (a) Design cruising speed, VC. For VC the following apply:
    (1) Where W/S'=wing loading at the design maximum takeoff weight, 
Vc (in knots) may not be less than--
    (i) 33 [radic](W/S) (for normal, utility, and commuter category 
airplanes);
    (ii) 36 [radic](W/S) (for acrobatic category airplanes).
    (2) For values of W/S more than 20, the multiplying factors may be 
decreased linearly with W/S to a value of 28.6 where W/S=100.
    (3) VC need not be more than 0.9 VH at sea level.
    (4) At altitudes where an MD is established, a cruising speed MC 
limited by compressibility may be selected.
    (b) Design dive speed VD. For VD, the following apply:
    (1) VD/MD may not be less than 1.25 VC/MC; and
    (2) With VC min, the required minimum design cruising speed, VD (in 
knots) may not be less than--
    (i) 1.40 Vc min (for normal and commuter category airplanes);
    (ii) 1.50 VC min (for utility category airplanes); and
    (iii) 1.55 VC min (for acrobatic category airplanes).
    (3) For values of W/S more than 20, the multiplying factors in 
paragraph (b)(2) of this section may be decreased linearly with W/S to a 
value of 1.35 where W/S=100.
    (4) Compliance with paragraphs (b)(1) and (2) of this section need 
not be shown if VD/MD is selected so that the minimum speed margin 
between VC/MC and VD/MD is the greater of the following:
    (i) The speed increase resulting when, from the initial condition of 
stabilized flight at VC/MC, the airplane is assumed to be upset, flown 
for 20 seconds along a flight path 7.5[deg] below the initial path, and 
then pulled up with a load factor of 1.5 (0.5 g. acceleration 
increment). At least 75 percent maximum continuous power for 
reciprocating engines, and maximum cruising power for turbines, or, if 
less, the power required for VC/MC for both kinds of engines, must be 
assumed until the pullup is initiated, at which point power reduction 
and pilot-controlled drag devices may be used; and either--
    (ii) Mach 0.05 for normal, utility, and acrobatic category airplanes 
(at altitudes where MD is established); or
    (iii) Mach 0.07 for commuter category airplanes (at altitudes where 
MD is established) unless a rational analysis, including the 
effects of automatic systems, is used to determine a lower margin. If a 
rational analysis is used, the minimum speed margin must be enough to 
provide for atmospheric variations (such as horizontal gusts), and the 
penetration of jet streams or cold fronts), instrument errors, airframe 
production variations, and must not be less than Mach 0.05.
    (c) Design maneuvering speed VA. For VA, the following applies:
    (1) VA may not be less than VS[radic]n where--
    (i) VS is a computed stalling speed with flaps retracted at the 
design weight, normally based on the maximum airplane normal force 
coefficients, CNA; and
    (ii) n is the limit maneuvering load factor used in design
    (2) The value of VA need not exceed the value of VC used in design.
    (d) Design speed for maximum gust intensity, VB. For VB, the 
following apply:
    (1) VB may not be less than the speed determined by the 
intersection of the line representing the maximum positive lift, 
CNMAX, and the line representing the rough air gust velocity 
on the gust V-n diagram, or VS1[radic] ng, 
whichever is less, where:
    (i) ng the positive airplane gust load factor due to 
gust, at speed VC (in accordance with Sec. 23.341), and at the 
particular weight under consideration; and
    (ii) VS1 is the stalling speed with the flaps retracted 
at the particular weight under consideration.
    (2) VB need not be greater than VC.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13088, Aug. 13, 1969; Amdt. 23-16, 40 FR 2577, Jan. 14, 1975; Amdt. 
23-34, 52 FR 1829, Jan. 15, 1987; Amdt. 23-24, 52 FR 34745, Sept. 14, 
1987; Amdt. 23-48, 61 FR 5143, Feb. 9, 1996]

[[Page 213]]



Sec. 23.337  Limit maneuvering load factors.

    (a) The positive limit maneuvering load factor n may not be less 
than--
    (1) 2.1+(24,000/(W+10,000)) for normal and commuter category 
airplanes, where W=design maximum takeoff weight, except that n need not 
be more than 3.8;
    (2) 4.4 for utility category airplanes; or
    (3) 6.0 for acrobatic category airplanes.
    (b) The negative limit maneuvering load factor may not be less 
than--
    (1) 0.4 times the positive load factor for the normal utility and 
commuter categories; or
    (2) 0.5 times the positive load factor for the acrobatic category.
    (c) Maneuvering load factors lower than those specified in this 
section may be used if the airplane has design features that make it 
impossible to exceed these values in flight.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13088, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987; Amdt. 
23-48, 61 FR 5144, Feb. 9, 1996]



Sec. 23.341  Gust loads factors.

    (a) Each airplane must be designed to withstand loads on each 
lifting surface resulting from gusts specified in Sec. 23.333(c).
    (b) The gust load for a canard or tandem wing configuration must be 
computed using a rational analysis, or may be computed in accordance 
with paragraph (c) of this section, provided that the resulting net 
loads are shown to be conservative with respect to the gust criteria of 
Sec. 23.333(c).
    (c) In the absence of a more rational analysis, the gust load 
factors must be computed as follows--
[GRAPHIC] [TIFF OMITTED] TR09FE96.010

Where--

Kg=0.88[micro]g/5.3+[micro]g=gust alleviation 
          factor;
[micro]g=2(W/S)/[rho] Cag=airplane mass ratio;
Ude=Derived gust velocities referred to in Sec. 23.333(c) (f.p.s.);
[rho]=Density of air (slugs/cu.ft.);
W/S=Wing loading (p.s.f.) due to the applicable weight of the airplane 
          in the particular load case.
W/S=Wing loading (p.s.f.);
C=Mean geometric chord (ft.);
g=Acceleration due to gravity (ft./sec.\2\)
V=Airplane equivalent speed (knots); and
a=Slope of the airplane normal force coefficient curve CNA per radian if 
          the gust loads are applied to the wings and horizontal tail 
          surfaces simultaneously by a rational method. The wing lift 
          curve slope CL per radian may be used when the gust load is 
          applied to the wings only and the horizontal tail gust loads 
          are treated as a separate condition.

[Amdt. 23-7, 34 FR 13088, Aug. 13, 1969, as amended by Amdt. 23-42, 56 
FR 352, Jan. 3, 1991; Amdt. 23-48, 61 FR 5144, Feb. 9, 1996]



Sec. 23.343  Design fuel loads.

    (a) The disposable load combinations must include each fuel load in 
the range from zero fuel to the selected maximum fuel load.
    (b) If fuel is carried in the wings, the maximum allowable weight of 
the airplane without any fuel in the wing tank(s) must be established as 
``maximum zero wing fuel weight,'' if it is less than the maximum 
weight.
    (c) For commuter category airplanes, a structural reserve fuel 
condition, not exceeding fuel necessary for 45 minutes of operation at 
maximum continuous power, may be selected. If a structural reserve fuel 
condition is selected, it must be used as the minimum fuel weight 
condition for showing compliance with the flight load requirements 
prescribed in this part and--
    (1) The structure must be designed to withstand a condition of zero 
fuel in the wing at limit loads corresponding to:
    (i) Ninety percent of the maneuvering load factors defined in Sec. 
23.337, and
    (ii) Gust velocities equal to 85 percent of the values prescribed in 
Sec. 23.333(c).
    (2) The fatigue evaluation of the structure must account for any 
increase in operating stresses resulting from the design condition of 
paragraph (c)(1) of this section.
    (3) The flutter, deformation, and vibration requirements must also 
be met with zero fuel in the wings.

[Doc. No. 27805, 61 FR 5144, Feb. 9, 1996]

[[Page 214]]



Sec. 23.345  High lift devices.

    (a) If flaps or similar high lift devices are to be used for 
takeoff, approach or landing, the airplane, with the flaps fully 
extended at VF, is assumed to be subjected to symmetrical 
maneuvers and gusts within the range determined by--
    (1) Maneuvering, to a positive limit load factor of 2.0; and
    (2) Positive and negative gust of 25 feet per second acting normal 
to the flight path in level flight.
    (b) VF must be assumed to be not less than 1.4 
VS or 1.8 VSF, whichever is greater, where--
    (1) VS is the computed stalling speed with flaps 
retracted at the design weight; and
    (2) VSF is the computed stalling speed with flaps fully 
extended at the design weight.
    (3) If an automatic flap load limiting device is used, the airplane 
may be designed for the critical combinations of airspeed and flap 
position allowed by that device.
    (c) In determining external loads on the airplane as a whole, 
thrust, slipstream, and pitching acceleration may be assumed to be zero.
    (d) The flaps, their operating mechanism, and their supporting 
structures, must be designed to withstand the conditions prescribed in 
paragraph (a) of this section. In addition, with the flaps fully 
extended at VF, the following conditions, taken separately, 
must be accounted for:
    (1) A head-on gust having a velocity of 25 feet per second (EAS), 
combined with propeller slipstream corresponding to 75 percent of 
maximum continuous power; and
    (2) The effects of propeller slipstream corresponding to maximum 
takeoff power.

[Doc. No. 27805, 61 FR 5144, Feb. 9, 1996]



Sec. 23.347  Unsymmetrical flight conditions.

    (a) The airplane is assumed to be subjected to the unsymmetrical 
flight conditions of Sec. Sec. 23.349 and 23.351. Unbalanced 
aerodynamic moments about the center of gravity must be reacted in a 
rational or conservative manner, considering the principal masses 
furnishing the reacting inertia forces.
    (b) Acrobatic category airplanes certified for flick maneuvers (snap 
roll) must be designed for additional asymmetric loads acting on the 
wing and the horizontal tail.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-48, 
61 FR 5144, Feb. 9, 1996]



Sec. 23.349  Rolling conditions.

    The wing and wing bracing must be designed for the following loading 
conditions:
    (a) Unsymmetrical wing loads appropriate to the category. Unless the 
following values result in unrealistic loads, the rolling accelerations 
may be obtained by modifying the symmetrical flight conditions in Sec. 
23.333(d) as follows:
    (1) For the acrobatic category, in conditions A and F, assume that 
100 percent of the semispan wing airload acts on one side of the plane 
of symmetry and 60 percent of this load acts on the other side.
    (2) For normal, utility, and commuter categories, in Condition A, 
assume that 100 percent of the semispan wing airload acts on one side of 
the airplane and 75 percent of this load acts on the other side.
    (b) The loads resulting from the aileron deflections and speeds 
specified in Sec. 23.455, in combination with an airplane load factor 
of at least two thirds of the positive maneuvering load factor used for 
design. Unless the following values result in unrealistic loads, the 
effect of aileron displacement on wing torsion may be accounted for by 
adding the following increment to the basic airfoil moment coefficient 
over the aileron portion of the span in the critical condition 
determined in Sec. 23.333(d):

[Delta]cm=-0.01[delta]

where--

[Delta]cm is the moment coefficient increment; and
[delta] is the down aileron deflection in degrees in the critical 
          condition.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13088, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987; Amdt. 
23-48, 61 FR 5144, Feb. 9, 1996]

[[Page 215]]



Sec. 23.351  Yawing conditions.

    The airplane must be designed for yawing loads on the vertical 
surfaces resulting from the loads specified in Sec. Sec. 23.441 through 
23.445.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-42, 56 FR 352, Jan. 3, 1991]



Sec. 23.361  Engine torque.

    (a) Each engine mount and its supporting structure must be designed 
for the effects of--
    (1) A limit engine torque corresponding to takeoff power and 
propeller speed acting simultaneously with 75 percent of the limit loads 
from flight condition A of Sec. 23.333(d);
    (2) A limit engine torque corresponding to maximum continuous power 
and propeller speed acting simultaneously with the limit loads from 
flight condition A of Sec. 23.333(d); and
    (3) For turbopropeller installations, in addition to the conditions 
specified in paragraphs (a)(1) and (a)(2) of this section, a limit 
engine torque corresponding to takeoff power and propeller speed, 
multiplied by a factor accounting for propeller control system 
malfunction, including quick feathering, acting simultaneously with lg 
level flight loads. In the absence of a rational analysis, a factor of 
1.6 must be used.
    (b) For turbine engine installations, the engine mounts and 
supporting structure must be designed to withstand each of the 
following:
    (1) A limit engine torque load imposed by sudden engine stoppage due 
to malfunction or structural failure (such as compressor jamming).
    (2) A limit engine torque load imposed by the maximum acceleration 
of the engine.
    (c) The limit engine torque to be considered under paragraph (a) of 
this section must be obtained by multiplying the mean torque by a factor 
of--
    (1) 1.25 for turbopropeller installations;
    (2) 1.33 for engines with five or more cylinders; and
    (3) Two, three, or four, for engines with four, three, or two 
cylinders, respectively.

[Amdt. 23-26, 45 FR 60171, Sept. 11, 1980, as amended by Amdt. 23-45, 58 
FR 42160, Aug. 6, 1993]



Sec. 23.363  Side load on engine mount.

    (a) Each engine mount and its supporting structure must be designed 
for a limit load factor in a lateral direction, for the side load on the 
engine mount, of not less than--
    (1) 1.33, or
    (2) One-third of the limit load factor for flight condition A.
    (b) The side load prescribed in paragraph (a) of this section may be 
assumed to be independent of other flight conditions.



Sec. 23.365  Pressurized cabin loads.

    For each pressurized compartment, the following apply:
    (a) The airplane structure must be strong enough to withstand the 
flight loads combined with pressure differential loads from zero up to 
the maximum relief valve setting.
    (b) The external pressure distribution in flight, and any stress 
concentrations, must be accounted for.
    (c) If landings may be made with the cabin pressurized, landing 
loads must be combined with pressure differential loads from zero up to 
the maximum allowed during landing.
    (d) The airplane structure must be strong enough to withstand the 
pressure differential loads corresponding to the maximum relief valve 
setting multiplied by a factor of 1.33, omitting other loads.
    (e) If a pressurized cabin has two or more compartments separated by 
bulkheads or a floor, the primary structure must be designed for the 
effects of sudden release of pressure in any compartment with external 
doors or windows. This condition must be investigated for the effects of 
failure of the largest opening in the compartment. The effects of 
intercompartmental venting may be considered.



Sec. 23.367  Unsymmetrical loads due to engine failure.

    (a) Turbopropeller airplanes must be designed for the unsymmetrical 
loads

[[Page 216]]

resulting from the failure of the critical engine including the 
following conditions in combination with a single malfunction of the 
propeller drag limiting system, considering the probable pilot 
corrective action on the flight controls:
    (1) At speeds between VMC and VD, the loads resulting from power 
failure because of fuel flow interruption are considered to be limit 
loads.
    (2) At speeds between VMC and VC, the loads resulting from the 
disconnection of the engine compressor from the turbine or from loss of 
the turbine blades are considered to be ultimate loads.
    (3) The time history of the thrust decay and drag buildup occurring 
as a result of the prescribed engine failures must be substantiated by 
test or other data applicable to the particular engine-propeller 
combination.
    (4) The timing and magnitude of the probable pilot corrective action 
must be conservatively estimated, considering the characteristics of the 
particular engine-propeller-airplane combination.
    (b) Pilot corrective action may be assumed to be initiated at the 
time maximum yawing velocity is reached, but not earlier than 2 seconds 
after the engine failure. The magnitude of the corrective action may be 
based on the limit pilot forces specified in Sec. 23.397 except that 
lower forces may be assumed where it is shown by analysis or test that 
these forces can control the yaw and roll resulting from the prescribed 
engine failure conditions.

[Amdt. 23-7, 34 FR 13089, Aug. 13, 1969]



Sec. 23.369  Rear lift truss.

    (a) If a rear lift truss is used, it must be designed to withstand 
conditions of reversed airflow at a design speed of--
    V=8.7 [radic](W/S) + 8.7 (knots), where W/S=wing loading at design 
maximum takeoff weight.
    (b) Either aerodynamic data for the particular wing section used, or 
a value of CL equalling -0.8 with a chordwise distribution that is 
triangular between a peak at the trailing edge and zero at the leading 
edge, must be used.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13089, Aug. 13, 1969; 34 FR 17509, Oct. 30, 1969; Amdt. 23-45, 58 FR 
42160, Aug. 6, 1993; Amdt. 23-48, 61 FR 5145, Feb. 9, 1996]



Sec. 23.371  Gyroscopic and aerodynamic loads.

    (a) Each engine mount and its supporting structure must be designed 
for the gyroscopic, inertial, and aerodynamic loads that result, with 
the engine(s) and propeller(s), if applicable, at maximum continuous 
r.p.m., under either:
    (1) The conditions prescribed in Sec. 23.351 and Sec. 23.423; or
    (2) All possible combinations of the following--
    (i) A yaw velocity of 2.5 radians per second;
    (ii) A pitch velocity of 1.0 radian per second;
    (iii) A normal load factor of 2.5; and
    (iv) Maximum continuous thrust.
    (b) For airplanes approved for aerobatic maneuvers, each engine 
mount and its supporting structure must meet the requirements of 
paragraph (a) of this section and be designed to withstand the load 
factors expected during combined maximum yaw and pitch velocities.
    (c) For airplanes certificated in the commuter category, each engine 
mount and its supporting structure must meet the requirements of 
paragraph (a) of this section and the gust conditions specified in Sec. 
23.341 of this part.

[Doc. No. 27805, 61 FR 5145, Feb. 9, 1996]



Sec. 23.373  Speed control devices.

    If speed control devices (such as spoilers and drag flaps) are 
incorporated for use in enroute conditions--
    (a) The airplane must be designed for the symmetrical maneuvers and 
gusts prescribed in Sec. Sec. 23.333, 23.337, and 23.341, and the 
yawing maneuvers and lateral gusts in Sec. Sec. 23.441 and 23.443, with 
the device extended at speeds up to the placard device extended speed; 
and
    (b) If the device has automatic operating or load limiting features, 
the airplane must be designed for the maneuver and gust conditions 
prescribed in

[[Page 217]]

paragraph (a) of this section at the speeds and corresponding device 
positions that the mechanism allows.

[Amdt. 23-7, 34 FR 13089, Aug. 13, 1969]

                    Control Surface and System Loads



Sec. 23.391  Control surface loads.

    The control surface loads specified in Sec. Sec. 23.397 through 
23.459 are assumed to occur in the conditions described in Sec. Sec. 
23.331 through 23.351.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-48, 
61 FR 5145, Feb. 9, 1996]



Sec. 23.393  Loads parallel to hinge line.

    (a) Control surfaces and supporting hinge brackets must be designed 
to withstand inertial loads acting parallel to the hinge line.
    (b) In the absence of more rational data, the inertial loads may be 
assumed to be equal to KW, where--
    (1) K=24 for vertical surfaces;
    (2) K=12 for horizontal surfaces; and
    (3) W=weight of the movable surfaces.

[Doc. No. 27805, 61 FR 5145, Feb. 9, 1996]



Sec. 23.395  Control system loads.

    (a) Each flight control system and its supporting structure must be 
designed for loads corresponding to at least 125 percent of the computed 
hinge moments of the movable control surface in the conditions 
prescribed in Sec. Sec. 23.391 through 23.459. In addition, the 
following apply:
    (1) The system limit loads need not exceed the higher of the loads 
that can be produced by the pilot and automatic devices operating the 
controls. However, autopilot forces need not be added to pilot forces. 
The system must be designed for the maximum effort of the pilot or 
autopilot, whichever is higher. In addition, if the pilot and the 
autopilot act in opposition, the part of the system between them may be 
designed for the maximum effort of the one that imposes the lesser load. 
Pilot forces used for design need not exceed the maximum forces 
prescribed in Sec. 23.397(b).
    (2) The design must, in any case, provide a rugged system for 
service use, considering jamming, ground gusts, taxiing downwind, 
control inertia, and friction. Compliance with this subparagraph may be 
shown by designing for loads resulting from application of the minimum 
forces prescribed in Sec. 23.397(b).
    (b) A 125 percent factor on computed hinge moments must be used to 
design elevator, aileron, and rudder systems. However, a factor as low 
as 1.0 may be used if hinge moments are based on accurate flight test 
data, the exact reduction depending upon the accuracy and reliability of 
the data.
    (c) Pilot forces used for design are assumed to act at the 
appropriate control grips or pads as they would in flight, and to react 
at the attachments of the control system to the control surface horns.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13089, Aug. 13, 1969]



Sec. 23.397  Limit control forces and torques.

    (a) In the control surface flight loading condition, the airloads on 
movable surfaces and the corresponding deflections need not exceed those 
that would result in flight from the application of any pilot force 
within the ranges specified in paragraph (b) of this section. In 
applying this criterion, the effects of control system boost and servo-
mechanisms, and the effects of tabs must be considered. The automatic 
pilot effort must be used for design if it alone can produce higher 
control surface loads than the human pilot.
    (b) The limit pilot forces and torques are as follows:

------------------------------------------------------------------------
                                   Maximum forces or
                                  torques for design
                                    weight, weight     Minimum forces or
             Control               equal to or less       torques \2\
                                   than 5,000 pounds
                                          \1\
------------------------------------------------------------------------
Aileron:
  Stick.........................  67 lbs............  40 lbs.
  Wheel \3\.....................  50 D in.-lbs \4\..  40 D in.-lbs.\4\
Elevator:
  Stick.........................  167 lbs...........  100 lbs.
  Wheel (symmetrical)...........  200 lbs...........  100 lbs.
  Wheel (unsymmetrical) \5\.....  ..................  100 lbs.
Rudder..........................  200 lbs...........  150 lbs.
------------------------------------------------------------------------
\1\ For design weight (W) more than 5,000 pounds, the specified maximum
  values must be increased linearly with weight to 1.18 times the
  specified values at a design weight of 12,500 pounds and for commuter
  category airplanes, the specified values must be increased linearly
  with weight to 1.35 times the specified values at a design weight of
  19,000 pounds.

[[Page 218]]

 
\2\ If the design of any individual set of control systems or surfaces
  makes these specified minimum forces or torques inapplicable, values
  corresponding to the present hinge moments obtained under Sec.
  23.415, but not less than 0.6 of the specified minimum forces or
  torques, may be used.
\3\ The critical parts of the aileron control system must also be
  designed for a single tangential force with a limit value of 1.25
  times the couple force determined from the above criteria.
\4\ D=wheel diameter (inches).
\5\ The unsymmetrical force must be applied at one of the normal
  handgrip points on the control wheel.


[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13089, Aug. 13, 1969; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. 
23-34, 52 FR 1829, Jan. 15, 1987; Amdt. 23-45, 58 FR 42160, Aug. 6, 
1993]



Sec. 23.399  Dual control system.

    (a) Each dual control system must be designed to withstand the force 
of the pilots operating in opposition, using individual pilot forces not 
less than the greater of--
    (1) 0.75 times those obtained under Sec. 23.395; or
    (2) The minimum forces specified in Sec. 23.397(b).
    (b) Each dual control system must be designed to withstand the force 
of the pilots applied together, in the same direction, using individual 
pilot forces not less than 0.75 times those obtained under Sec. 23.395.

[Doc. No. 27805, 61 FR 5145, Feb. 9, 1996]



Sec. 23.405  Secondary control system.

    Secondary controls, such as wheel brakes, spoilers, and tab 
controls, must be designed for the maximum forces that a pilot is likely 
to apply to those controls.



Sec. 23.407  Trim tab effects.

    The effects of trim tabs on the control surface design conditions 
must be accounted for only where the surface loads are limited by 
maximum pilot effort. In these cases, the tabs are considered to be 
deflected in the direction that would assist the pilot. These 
deflections must correspond to the maximum degree of ``out of trim'' 
expected at the speed for the condition under consideration.



Sec. 23.409  Tabs.

    Control surface tabs must be designed for the most severe 
combination of airspeed and tab deflection likely to be obtained within 
the flight envelope for any usable loading condition.



Sec. 23.415  Ground gust conditions.

    (a) The control system must be investigated as follows for control 
surface loads due to ground gusts and taxiing downwind:
    (1) If an investigation of the control system for ground gust loads 
is not required by paragraph (a)(2) of this section, but the applicant 
elects to design a part of the control system of these loads, these 
loads need only be carried from control surface horns through the 
nearest stops or gust locks and their supporting structures.
    (2) If pilot forces less than the minimums specified in Sec. 
23.397(b) are used for design, the effects of surface loads due to 
ground gusts and taxiing downwind must be investigated for the entire 
control system according to the formula:

H=K c S q

where--

H=limit hinge moment (ft.-lbs.);
c=mean chord of the control surface aft of the hinge line (ft.);
S=area of control surface aft of the hinge line (sq. ft.);
q=dynamic pressure (p.s.f.) based on a design speed not less than 14.6 
          [radic](W/S) + 14.6 (f.p.s.) where W/S=wing loading at design 
          maximum weight, except that the design speed need not exceed 
          88 (f.p.s.);
K=limit hinge moment factor for ground gusts derived in paragraph (b) of 
          this section. (For ailerons and elevators, a positive value of 
          K indicates a moment tending to depress the surface and a 
          negative value of K indicates a moment tending to raise the 
          surface).

    (b) The limit hinge moment factor K for ground gusts must be derived 
as follows:

------------------------------------------------------------------------
            Surface                 K          Position of controls
------------------------------------------------------------------------
(a) Aileron....................     0.75  Control column locked lashed
                                           in mid-position.
(b) Aileron....................   
Sec. 23.421  Balancing loads.

    (a) A horizontal surface balancing load is a load necessary to 
maintain equilibrium in any specified flight condition with no pitching 
acceleration.
    (b) Horizontal balancing surfaces must be designed for the balancing 
loads occurring at any point on the limit maneuvering envelope and in 
the flap conditions specified in Sec. 23.345.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13089, Aug. 13, 1969; Amdt. 23-42, 56 FR 352, Jan. 3, 1991]



Sec. 23.423  Maneuvering loads.

    Each horizontal surface and its supporting structure, and the main 
wing of a canard or tandem wing configuration, if that surface has pitch 
control, must be designed for the maneuvering loads imposed by the 
following conditions:
    (a) A sudden movement of the pitching control, at the speed 
VA, to the maximum aft movement, and the maximum forward 
movement, as limited by the control stops, or pilot effort, whichever is 
critical.
    (b) A sudden aft movement of the pitching control at speeds above 
VA, followed by a forward movement of the pitching control 
resulting in the following combinations of normal and angular 
acceleration:

------------------------------------------------------------------------
                                     Normal
           Condition              acceleration     Angular acceleration
                                       (n)            (radian/sec2)
------------------------------------------------------------------------
Nose-up pitching...............  1.0             +39nm/Vx(nm-1.5)
Nose-down pitching.............  nm              -39nm/Vx(nm-1.5)
------------------------------------------------------------------------


where--
    (1) nm=positive limit maneuvering load factor used in the 
design of the airplane; and
    (2) V=initial speed in knots.
    The conditions in this paragraph involve loads corresponding to the 
loads that may occur in a ``checked maneuver'' (a maneuver in which the 
pitching control is suddenly displaced in one direction and then 
suddenly moved in the opposite direction). The deflections and timing of 
the ``checked maneuver'' must avoid exceeding the limit maneuvering load 
factor. The total horizontal surface load for both nose-up and nose-down 
pitching conditions is the sum of the balancing loads at V and the 
specified value of the normal load factor n, plus the maneuvering load 
increment due to the specified value of the angular acceleration.

[Amdt. 23-42, 56 FR 353, Jan. 3, 1991; 56 FR 5455, Feb. 11, 1991]



Sec. 23.425  Gust loads.

    (a) Each horizontal surface, other than a main wing, must be 
designed for loads resulting from--
    (1) Gust velocities specified in Sec. 23.333(c) with flaps 
retracted; and
    (2) Positive and negative gusts of 25 f.p.s. nominal intensity at VF 
corresponding to the flight conditions specified in Sec. 23.345(a)(2).
    (b) [Reserved]
    (c) When determining the total load on the horizontal surfaces for 
the conditions specified in paragraph (a) of this section, the initial 
balancing loads for steady unaccelerated flight at the pertinent design 
speeds VF, VC, and VD must first be 
determined. The incremental load resulting from the gusts must be added 
to the initial balancing load to obtain the total load.
    (d) In the absence of a more rational analysis, the incremental load 
due to the gust must be computed as follows only on airplane 
configurations with aft-mounted, horizontal surfaces, unless its use 
elsewhere is shown to be conservative:
[GRAPHIC] [TIFF OMITTED] TC28SE91.002

where--

[Delta]Lht=Incremental horizontal tailload (lbs.);
Kg=Gust alleviation factor defined in Sec. 23.341;
Ude=Derived gust velocity (f.p.s.);
V=Airplane equivalent speed (knots);
aht=Slope of aft horizontal lift curve (per radian)

[[Page 220]]

Sht=Area of aft horizontal lift surface (ft\2\); and
[GRAPHIC] [TIFF OMITTED] TC28SE91.003


[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13089 Aug. 13, 1969; Amdt. 23-42, 56 FR 353, Jan. 3, 1991]



Sec. 23.427  Unsymmetrical loads.

    (a) Horizontal surfaces other than main wing and their supporting 
structure must be designed for unsymmetrical loads arising from yawing 
and slipstream effects, in combination with the loads prescribed for the 
flight conditions set forth in Sec. Sec. 23.421 through 23.425.
    (b) In the absence of more rational data for airplanes that are 
conventional in regard to location of engines, wings, horizontal 
surfaces other than main wing, and fuselage shape:
    (1) 100 percent of the maximum loading from the symmetrical flight 
conditions may be assumed on the surface on one side of the plane of 
symmetry; and
    (2) The following percentage of that loading must be applied to the 
opposite side:

    Percent=100-10 (n-1), where n is the specified positive maneuvering 
load factor, but this value may not be more than 80 percent.

    (c) For airplanes that are not conventional (such as airplanes with 
horizontal surfaces other than main wing having appreciable dihedral or 
supported by the vertical tail surfaces) the surfaces and supporting 
structures must be designed for combined vertical and horizontal surface 
loads resulting from each prescribed flight condition taken separately.

[Amdt. 23-14, 38 FR 31820, Nov. 19, 1973, as amended by Amdt. 23-42, 56 
FR 353, Jan. 3, 1991]

                            Vertical Surfaces



Sec. 23.441  Maneuvering loads.

    (a) At speeds up to VA, the vertical surfaces must be designed to 
withstand the following conditions. In computing the loads, the yawing 
velocity may be assumed to be zero:
    (1) With the airplane in unaccelerated flight at zero yaw, it is 
assumed that the rudder control is suddenly displaced to the maximum 
deflection, as limited by the control stops or by limit pilot forces.
    (2) With the rudder deflected as specified in paragraph (a)(1) of 
this section, it is assumed that the airplane yaws to the overswing 
sideslip angle. In lieu of a rational analysis, an overswing angle equal 
to 1.5 times the static sideslip angle of paragraph (a)(3) of this 
section may be assumed.
    (3) A yaw angle of 15 degrees with the rudder control maintained in 
the neutral position (except as limited by pilot strength).
    (b) For commuter category airplanes, the loads imposed by the 
following additional maneuver must be substantiated at speeds from 
VA to VD/MD. When computing the tail 
loads--
    (1) The airplane must be yawed to the largest attainable steady 
state sideslip angle, with the rudder at maximum deflection caused by 
any one of the following:
    (i) Control surface stops;
    (ii) Maximum available booster effort;
    (iii) Maximum pilot rudder force as shown below:

[[Page 221]]

[GRAPHIC] [TIFF OMITTED] TR09FE96.006

    (2) The rudder must be suddenly displaced from the maximum 
deflection to the neutral position.
    (c) The yaw angles specified in paragraph (a)(3) of this section may 
be reduced if the yaw angle chosen for a particular speed cannot be 
exceeded in--

[[Page 222]]

    (1) Steady slip conditions;
    (2) Uncoordinated rolls from steep banks; or
    (3) Sudden failure of the critical engine with delayed corrective 
action.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13090, Aug. 13, 1969; Amdt. 23-14, 38 FR 31821, Nov. 19, 1973; Amdt. 
23-28, 47 FR 13315, Mar. 29, 1982; Amdt. 23-42, 56 FR 353, Jan. 3, 1991; 
Amdt. 23-48, 61 FR 5145, Feb. 9, 1996]



Sec. 23.443  Gust loads.

    (a) Vertical surfaces must be designed to withstand, in 
unaccelerated flight at speed VC, lateral gusts of the values prescribed 
for VC in Sec. 23.333(c).
    (b) In addition, for commuter category airplanes, the airplane is 
assumed to encounter derived gusts normal to the plane of symmetry while 
in unaccelerated flight at VB, VC, VD, and VF. The derived gusts and 
airplane speeds corresponding to these conditions, as determined by 
Sec. Sec. 23.341 and 23.345, must be investigated. The shape of the 
gust must be as specified in Sec. 23.333(c)(2)(i).
    (c) In the absence of a more rational analysis, the gust load must 
be computed as follows:
[GRAPHIC] [TIFF OMITTED] TR09FE96.000

Where--

Lvt=Vertical surface loads (lbs.);
[GRAPHIC] [TIFF OMITTED] TR09FE96.001

[GRAPHIC] [TIFF OMITTED] TR09FE96.002

Ude=Derived gust velocity (f.p.s.);
[rho]=Air density (slugs/cu.ft.);
W=the applicable weight of the airplane in the particular load case 
          (lbs.);
Svt=Area of vertical surface (ft.\2\);
ct=Mean geometric chord of vertical surface (ft.);
avt=Lift curve slope of vertical surface (per radian);
K=Radius of gyration in yaw (ft.);
lvt=Distance from airplane c.g. to lift center of vertical 
          surface (ft.);
g=Acceleration due to gravity (ft./sec.\2\); and
V=Equivalent airspeed (knots).

[Amdt. 23-7, 34 FR 13090, Aug. 13, 1969, as amended by Amdt. 23-34, 52 
FR 1830, Jan. 15, 1987; 52 FR 7262, Mar. 9, 1987; Amdt. 23-24, 52 FR 
34745, Sept. 14, 1987; Amdt. 23-42, 56 FR 353, Jan. 3, 1991; Amdt. 23-
48, 61 FR 5147, Feb. 9, 1996]



Sec. 23.445  Outboard fins or winglets.

    (a) If outboard fins or winglets are included on the horizontal 
surfaces or wings, the horizontal surfaces or wings must be designed for 
their maximum load in combination with loads induced by the fins or 
winglets and moments or forces exerted on the horizontal surfaces or 
wings by the fins or winglets.
    (b) If outboard fins or winglets extend above and below the 
horizontal surface, the critical vertical surface loading (the load per 
unit area as determined under Sec. Sec. 23.441 and 23.443) must be 
applied to--
    (1) The part of the vertical surfaces above the horizontal surface 
with 80 percent of that loading applied to the part below the horizontal 
surface; and
    (2) The part of the vertical surfaces below the horizontal surface 
with 80 percent of that loading applied to the part above the horizontal 
surface.
    (c) The end plate effects of outboard fins or winglets must be taken 
into account in applying the yawing conditions of Sec. Sec. 23.441 and 
23.443 to the vertical surfaces in paragraph (b) of this section.
    (d) When rational methods are used for computing loads, the 
maneuvering loads of Sec. 23.441 on the vertical surfaces and the one-g 
horizontal surface load, including induced loads on the horizontal 
surface and moments or forces exerted on the horizontal surfaces by the 
vertical surfaces, must be applied simultaneously for the structural 
loading condition.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31821, Nov. 19, 1973; Amdt. 23-42, 56 FR 353, Jan. 3, 1991]

                      Ailerons and Special Devices



Sec. 23.455  Ailerons.

    (a) The ailerons must be designed for the loads to which they are 
subjected--
    (1) In the neutral position during symmetrical flight conditions; 
and

[[Page 223]]

    (2) By the following deflections (except as limited by pilot 
effort), during unsymmetrical flight conditions:
    (i) Sudden maximum displacement of the aileron control at VA. 
Suitable allowance may be made for control system deflections.
    (ii) Sufficient deflection at VC, where VC is more than VA, to 
produce a rate of roll not less than obtained in paragraph (a)(2)(i) of 
this section.
    (iii) Sufficient deflection at VD to produce a rate of roll not less 
than one-third of that obtained in paragraph (a)(2)(i) of this section.
    (b) [Reserved]

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13090, Aug. 13, 1969; Amdt. 23-42, 56 FR 353, Jan. 3, 1991]



Sec. 23.459  Special devices.

    The loading for special devices using aerodynamic surfaces (such as 
slots and spoilers) must be determined from test data.

                              Ground Loads



Sec. 23.471  General.

    The limit ground loads specified in this subpart are considered to 
be external loads and inertia forces that act upon an airplane 
structure. In each specified ground load condition, the external 
reactions must be placed in equilibrium with the linear and angular 
inertia forces in a rational or conservative manner.



Sec. 23.473  Ground load conditions and assumptions.

    (a) The ground load requirements of this subpart must be complied 
with at the design maximum weight except that Sec. Sec. 23.479, 23.481, 
and 23.483 may be complied with at a design landing weight (the highest 
weight for landing conditions at the maximum descent velocity) allowed 
under paragraphs (b) and (c) of this section.
    (b) The design landing weight may be as low as--
    (1) 95 percent of the maximum weight if the minimum fuel capacity is 
enough for at least one-half hour of operation at maximum continuous 
power plus a capacity equal to a fuel weight which is the difference 
between the design maximum weight and the design landing weight; or
    (2) The design maximum weight less the weight of 25 percent of the 
total fuel capacity.
    (c) The design landing weight of a multiengine airplane may be less 
than that allowed under paragraph (b) of this section if--
    (1) The airplane meets the one-engine-inoperative climb requirements 
of Sec. 23.67(b)(1) or (c); and
    (2) Compliance is shown with the fuel jettisoning system 
requirements of Sec. 23.1001.
    (d) The selected limit vertical inertia load factor at the center of 
gravity of the airplane for the ground load conditions prescribed in 
this subpart may not be less than that which would be obtained when 
landing with a descent velocity (V), in feet per second, equal to 4.4 
(W/S)\1/4\, except that this velocity need not be more than 10 feet per 
second and may not be less than seven feet per second.
    (e) Wing lift not exceeding two-thirds of the weight of the airplane 
may be assumed to exist throughout the landing impact and to act through 
the center of gravity. The ground reaction load factor may be equal to 
the inertia load factor minus the ratio of the above assumed wing lift 
to the airplane weight.
    (f) If energy absorption tests are made to determine the limit load 
factor corresponding to the required limit descent velocities, these 
tests must be made under Sec. 23.723(a).
    (g) No inertia load factor used for design purposes may be less than 
2.67, nor may the limit ground reaction load factor be less than 2.0 at 
design maximum weight, unless these lower values will not be exceeded in 
taxiing at speeds up to takeoff speed over terrain as rough as that 
expected in service.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13090, Aug. 13, 1969; Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. 
23-45, 58 FR 42160, Aug. 6, 1993; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996]



Sec. 23.477  Landing gear arrangement.

    Sections 23.479 through 23.483, or the conditions in appendix C, 
apply to airplanes with conventional arrangements

[[Page 224]]

of main and nose gear, or main and tail gear.



Sec. 23.479  Level landing conditions.

    (a) For a level landing, the airplane is assumed to be in the 
following attitudes:
    (1) For airplanes with tail wheels, a normal level flight attitude.
    (2) For airplanes with nose wheels, attitudes in which--
    (i) The nose and main wheels contact the ground simultaneously; and
    (ii) The main wheels contact the ground and the nose wheel is just 
clear of the ground.

The attitude used in paragraph (a)(2)(i) of this section may be used in 
the analysis required under paragraph (a)(2)(ii) of this section.
    (b) When investigating landing conditions, the drag components 
simulating the forces required to accelerate the tires and wheels up to 
the landing speed (spin-up) must be properly combined with the 
corresponding instantaneous vertical ground reactions, and the forward-
acting horizontal loads resulting from rapid reduction of the spin-up 
drag loads (spring-back) must be combined with vertical ground reactions 
at the instant of the peak forward load, assuming wing lift and a tire-
sliding coefficient of friction of 0.8. However, the drag loads may not 
be less than 25 percent of the maximum vertical ground reactions 
(neglecting wing lift).
    (c) In the absence of specific tests or a more rational analysis for 
determining the wheel spin-up and spring-back loads for landing 
conditions, the method set forth in appendix D of this part must be 
used. If appendix D of this part is used, the drag components used for 
design must not be less than those given by appendix C of this part.
    (d) For airplanes with tip tanks or large overhung masses (such as 
turbo-propeller or jet engines) supported by the wing, the tip tanks and 
the structure supporting the tanks or overhung masses must be designed 
for the effects of dynamic responses under the level landing conditions 
of either paragraph (a)(1) or (a)(2)(ii) of this section. In evaluating 
the effects of dynamic response, an airplane lift equal to the weight of 
the airplane may be assumed.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 
41 FR 55464, Dec. 20, 1976; Amdt. 23-45, 58 FR 42160, Aug. 6, 1993]



Sec. 23.481  Tail down landing conditions.

    (a) For a tail down landing, the airplane is assumed to be in the 
following attitudes:
    (1) For airplanes with tail wheels, an attitude in which the main 
and tail wheels contact the ground simultaneously.
    (2) For airplanes with nose wheels, a stalling attitude, or the 
maximum angle allowing ground clearance by each part of the airplane, 
whichever is less.
    (b) For airplanes with either tail or nose wheels, ground reactions 
are assumed to be vertical, with the wheels up to speed before the 
maximum vertical load is attained.



Sec. 23.483  One-wheel landing conditions.

    For the one-wheel landing condition, the airplane is assumed to be 
in the level attitude and to contact the ground on one side of the main 
landing gear. In this attitude, the ground reactions must be the same as 
those obtained on that side under Sec. 23.479.



Sec. 23.485  Side load conditions.

    (a) For the side load condition, the airplane is assumed to be in a 
level attitude with only the main wheels contacting the ground and with 
the shock absorbers and tires in their static positions.
    (b) The limit vertical load factor must be 1.33, with the vertical 
ground reaction divided equally between the main wheels.
    (c) The limit side inertia factor must be 0.83, with the side ground 
reaction divided between the main wheels so that--
    (1) 0.5 (W) is acting inboard on one side; and
    (2) 0.33 (W) is acting outboard on the other side.
    (d) The side loads prescribed in paragraph (c) of this section are 
assumed to be applied at the ground contact point

[[Page 225]]

and the drag loads may be assumed to be zero.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42160, Aug. 6, 1993]



Sec. 23.493  Braked roll conditions.

    Under braked roll conditions, with the shock absorbers and tires in 
their static positions, the following apply:
    (a) The limit vertical load factor must be 1.33.
    (b) The attitudes and ground contacts must be those described in 
Sec. 23.479 for level landings.
    (c) A drag reaction equal to the vertical reaction at the wheel 
multiplied by a coefficient of friction of 0.8 must be applied at the 
ground contact point of each wheel with brakes, except that the drag 
reaction need not exceed the maximum value based on limiting brake 
torque.



Sec. 23.497  Supplementary conditions for tail wheels.

    In determining the ground loads on the tail wheel and affected 
supporting structures, the following apply:
    (a) For the obstruction load, the limit ground reaction obtained in 
the tail down landing condition is assumed to act up and aft through the 
axle at 45 degrees. The shock absorber and tire may be assumed to be in 
their static positions.
    (b) For the side load, a limit vertical ground reaction equal to the 
static load on the tail wheel, in combination with a side component of 
equal magnitude, is assumed. In addition--
    (1) If a swivel is used, the tail wheel is assumed to be swiveled 90 
degrees to the airplane longitudinal axis with the resultant ground load 
passing through the axle;
    (2) If a lock, steering device, or shimmy damper is used, the tail 
wheel is also assumed to be in the trailing position with the side load 
acting at the ground contact point; and
    (3) The shock absorber and tire are assumed to be in their static 
positions.
    (c) If a tail wheel, bumper, or an energy absorption device is 
provided to show compliance with Sec. 23.925(b), the following apply:
    (1) Suitable design loads must be established for the tail wheel, 
bumper, or energy absorption device; and
    (2) The supporting structure of the tail wheel, bumper, or energy 
absorption device must be designed to withstand the loads established in 
paragraph (c)(1) of this section.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-48, 
61 FR 5147, Feb. 9, 1996]



Sec. 23.499  Supplementary conditions for nose wheels.

    In determining the ground loads on nose wheels and affected 
supporting structures, and assuming that the shock absorbers and tires 
are in their static positions, the following conditions must be met:
    (a) For aft loads, the limit force components at the axle must be--
    (1) A vertical component of 2.25 times the static load on the wheel; 
and
    (2) A drag component of 0.8 times the vertical load.
    (b) For forward loads, the limit force components at the axle must 
be--
    (1) A vertical component of 2.25 times the static load on the wheel; 
and
    (2) A forward component of 0.4 times the vertical load.
    (c) For side loads, the limit force components at ground contact 
must be--
    (1) A vertical component of 2.25 times the static load on the wheel; 
and
    (2) A side component of 0.7 times the vertical load.
    (d) For airplanes with a steerable nose wheel that is controlled by 
hydraulic or other power, at design takeoff weight with the nose wheel 
in any steerable position, the application of 1.33 times the full 
steering torque combined with a vertical reaction equal to 1.33 times 
the maximum static reaction on the nose gear must be assumed. However, 
if a torque limiting device is installed, the steering torque can be 
reduced to the maximum value allowed by that device.
    (e) For airplanes with a steerable nose wheel that has a direct 
mechanical connection to the rudder pedals, the mechanism must be 
designed to withstand the steering torque for the

[[Page 226]]

maximum pilot forces specified in Sec. 23.397(b).

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-48, 
61 FR 5147, Feb. 9, 1996]



Sec. 23.505  Supplementary conditions for skiplanes.

    In determining ground loads for skiplanes, and assuming that the 
airplane is resting on the ground with one main ski frozen at rest and 
the other skis free to slide, a limit side force equal to 0.036 times 
the design maximum weight must be applied near the tail assembly, with a 
factor of safety of 1.

[Amdt. 23-7, 34 FR 13090, Aug. 13, 1969]



Sec. 23.507  Jacking loads.

    (a) The airplane must be designed for the loads developed when the 
aircraft is supported on jacks at the design maximum weight assuming the 
following load factors for landing gear jacking points at a three-point 
attitude and for primary flight structure jacking points in the level 
attitude:
    (1) Vertical-load factor of 1.35 times the static reactions.
    (2) Fore, aft, and lateral load factors of 0.4 times the vertical 
static reactions.
    (b) The horizontal loads at the jack points must be reacted by 
inertia forces so as to result in no change in the direction of the 
resultant loads at the jack points.
    (c) The horizontal loads must be considered in all combinations with 
the vertical load.

[Amdt. 23-14, 38 FR 31821, Nov. 19, 1973]



Sec. 23.509  Towing loads.

    The towing loads of this section must be applied to the design of 
tow fittings and their immediate attaching structure.
    (a) The towing loads specified in paragraph (d) of this section must 
be considered separately. These loads must be applied at the towing 
fittings and must act parallel to the ground. In addition:
    (1) A vertical load factor equal to 1.0 must be considered acting at 
the center of gravity; and
    (2) The shock struts and tires must be in there static positions.
    (b) For towing points not on the landing gear but near the plane of 
symmetry of the airplane, the drag and side tow load components 
specified for the auxiliary gear apply. For towing points located 
outboard of the main gear, the drag and side tow load components 
specified for the main gear apply. Where the specified angle of swivel 
cannot be reached, the maximum obtainable angle must be used.
    (c) The towing loads specified in paragraph (d) of this section must 
be reacted as follows:
    (1) The side component of the towing load at the main gear must be 
reacted by a side force at the static ground line of the wheel to which 
the load is applied.
    (2) The towing loads at the auxiliary gear and the drag components 
of the towing loads at the main gear must be reacted as follows:
    (i) A reaction with a maximum value equal to the vertical reaction 
must be applied at the axle of the wheel to which the load is applied. 
Enough airplane inertia to achieve equilibrium must be applied.
    (ii) The loads must be reacted by airplane inertia.
    (d) The prescribed towing loads are as follows, where W is the 
design maximum weight:

----------------------------------------------------------------------------------------------------------------
                                                                                        Load
               Tow point                         Position         ----------------------------------------------
                                                                    Magnitude    No.            Direction
----------------------------------------------------------------------------------------------------------------
Main gear.............................  .........................      0.225W        1  Forward, parallel to
                                                                                     2   drag axis.
                                                                                     3  Forward, at 30[deg] to
                                                                                     4   drag axis.
                                                                                        Aft, parallel to drag
                                                                                         axis.
                                                                                        Aft, at 30[deg] to drag
                                                                                         axis.
----------------------------------------------------------------------------------------------------------------
Auxiliary gear........................  Swiveled forward.........        0.3W        5  Forward.
                                                                                     6  Aft.
                                        Swiveled aft.............        0.3W        7  Forward.
                                                                                     8  Aft.

[[Page 227]]

 
                                        Swiveled 45[deg] from           0.15W        9  Forward, in plane of
                                         forward.                                   10   wheel.
                                                                                        Aft, in plane of wheel.
                                        Swiveled 45[deg] from aft       0.15W       11  Forward, in plane of
                                                                                    12   wheel.
                                                                                        Aft, in plane of wheel.
----------------------------------------------------------------------------------------------------------------


[Amdt. 23-14, 38 FR 31821, Nov. 19, 1973]



Sec. 23.511  Ground load; unsymmetrical loads on multiple-wheel units.

    (a) Pivoting loads. The airplane is assumed to pivot about on side 
of the main gear with--
    (1) The brakes on the pivoting unit locked; and
    (2) Loads corresponding to a limit vertical load factor of 1, and 
coefficient of friction of 0.8 applied to the main gear and its 
supporting structure.
    (b) Unequal tire loads. The loads established under Sec. Sec. 
23.471 through 23.483 must be applied in turn, in a 60/40 percent 
distribution, to the dual wheels and tires in each dual wheel landing 
gear unit.
    (c) Deflated tire loads. For the deflated tire condition--
    (1) 60 percent of the loads established under Sec. Sec. 23.471 
through 23.483 must be applied in turn to each wheel in a landing gear 
unit; and
    (2) 60 percent of the limit drag and side loads, and 100 percent of 
the limit vertical load established under Sec. Sec. 23.485 and 23.493 
or lesser vertical load obtained under paragraph (c)(1) of this section, 
must be applied in turn to each wheel in the dual wheel landing gear 
unit.

[Amdt. 23-7, 34 FR 13090, Aug. 13, 1969]

                               Water Loads



Sec. 23.521  Water load conditions.

    (a) The structure of seaplanes and amphibians must be designed for 
water loads developed during takeoff and landing with the seaplane in 
any attitude likely to occur in normal operation at appropriate forward 
and sinking velocities under the most severe sea conditions likely to be 
encountered.
    (b) Unless the applicant makes a rational analysis of the water 
loads, Sec. Sec. 23.523 through 23.537 apply.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42160, Aug. 6, 1993; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996]



Sec. 23.523  Design weights and center of gravity positions.

    (a) Design weights. The water load requirements must be met at each 
operating weight up to the design landing weight except that, for the 
takeoff condition prescribed in Sec. 23.531, the design water takeoff 
weight (the maximum weight for water taxi and takeoff run) must be used.
    (b) Center of gravity positions. The critical centers of gravity 
within the limits for which certification is requested must be 
considered to reach maximum design loads for each part of the seaplane 
structure.

[Doc. No. 26269, 58 FR 42160, Aug. 6, 1993]



Sec. 23.525  Application of loads.

    (a) Unless otherwise prescribed, the seaplane as a whole is assumed 
to be subjected to the loads corresponding to the load factors specified 
in Sec. 23.527.
    (b) In applying the loads resulting from the load factors prescribed 
in Sec. 23.527, the loads may be distributed over the hull or main 
float bottom (in order to avoid excessive local shear loads and bending 
moments at the location of water load application) using pressures not 
less than those prescribed in Sec. 23.533(c).
    (c) For twin float seaplanes, each float must be treated as an 
equivalent hull on a fictitious seaplane with a weight equal to one-half 
the weight of the twin float seaplane.
    (d) Except in the takeoff condition of Sec. 23.531, the aerodynamic 
lift on the

[[Page 228]]

seaplane during the impact is assumed to be \2/3\ of the weight of the 
seaplane.

[Doc. No. 26269, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]



Sec. 23.527  Hull and main float load factors.

    (a) Water reaction load factors nw must be computed in 
the following manner:
    (1) For the step landing case
    [GRAPHIC] [TIFF OMITTED] TC28SE91.004
    
    (2) For the bow and stern landing cases
    [GRAPHIC] [TIFF OMITTED] TC28SE91.005
    
    (b) The following values are used:
    (1) nw=water reaction load factor (that is, the water 
reaction divided by seaplane weight).
    (2) C1=empirical seaplane operations factor equal to 
0.012 (except that this factor may not be less than that necessary to 
obtain the minimum value of step load factor of 2.33).
    (3) VSO=seaplane stalling speed in knots with flaps 
extended in the appropriate landing position and with no slipstream 
effect.
    (4) [beta]=Angle of dead rise at the longitudinal station at which 
the load factor is being determined in accordance with figure 1 of 
appendix I of this part.
    (5) W=seaplane landing weight in pounds.
    (6) K1=empirical hull station weighing factor, in 
accordance with figure 2 of appendix I of this part.
    (7) rx=ratio of distance, measured parallel to hull 
reference axis, from the center of gravity of the seaplane to the hull 
longitudinal station at which the load factor is being computed to the 
radius of gyration in pitch of the seaplane, the hull reference axis 
being a straight line, in the plane of symmetry, tangential to the keel 
at the main step.
    (c) For a twin float seaplane, because of the effect of flexibility 
of the attachment of the floats to the seaplane, the factor 
K1 may be reduced at the bow and stern to 0.8 of the value 
shown in figure 2 of appendix I of this part. This reduction applies 
only to the design of the carrythrough and seaplane structure.

[Doc. No. 26269, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]



Sec. 23.529  Hull and main float landing conditions.

    (a) Symmetrical step, bow, and stern landing. For symmetrical step, 
bow, and stern landings, the limit water reaction load factors are those 
computed under Sec. 23.527. In addition--
    (1) For symmetrical step landings, the resultant water load must be 
applied at the keel, through the center of gravity, and must be directed 
perpendicularly to the keel line;
    (2) For symmetrical bow landings, the resultant water load must be 
applied at the keel, one-fifth of the longitudinal distance from the bow 
to the step, and must be directed perpendicularly to the keel line; and
    (3) For symmetrical stern landings, the resultant water load must be 
applied at the keel, at a point 85 percent of the longitudinal distance 
from the step to the stern post, and must be directed perpendicularly to 
the keel line.
    (b) Unsymmetrical landing for hull and single float seaplanes. 
Unsymmetrical step, bow, and stern landing conditions must be 
investigated. In addition--
    (1) The loading for each condition consists of an upward component 
and a side component equal, respectively, to 0.75 and 0.25 tan [beta] 
times the resultant load in the corresponding symmetrical landing 
condition; and
    (2) The point of application and direction of the upward component 
of the load is the same as that in the symmetrical condition, and the 
point of application of the side component is at the same longitudinal 
station as the upward component but is directed inward perpendicularly 
to the plane of symmetry at a point midway between the keel and chine 
lines.
    (c) Unsymmetrical landing; twin float seaplanes. The unsymmetrical 
loading consists of an upward load at the step of each float of 0.75 and 
a side load of 0.25 tan [beta] at one float times the step landing load 
reached under Sec. 23.527. The

[[Page 229]]

side load is directed inboard, perpendicularly to the plane of symmetry 
midway between the keel and chine lines of the float, at the same 
longitudinal station as the upward load.

[Doc. No. 26269, 58 FR 42161, Aug. 6, 1993]



Sec. 23.531  Hull and main float takeoff condition.

    For the wing and its attachment to the hull or main float--
    (a) The aerodynamic wing lift is assumed to be zero; and
    (b) A downward inertia load, corresponding to a load factor computed 
from the following formula, must be applied:
[GRAPHIC] [TIFF OMITTED] TC28SE91.006

Where--

n=inertia load factor;
CTO=empirical seaplane operations factor equal to 0.004;
VS1=seaplane stalling speed (knots) at the design takeoff 
          weight with the flaps extended in the appropriate takeoff 
          position;
[beta]=angle of dead rise at the main step (degrees); and
W=design water takeoff weight in pounds.

[Doc. No. 26269, 58 FR 42161, Aug. 6, 1993]



Sec. 23.533  Hull and main float bottom pressures.

    (a) General. The hull and main float structure, including frames and 
bulkheads, stringers, and bottom plating, must be designed under this 
section.
    (b) Local pressures. For the design of the bottom plating and 
stringers and their attachments to the supporting structure, the 
following pressure distributions must be applied:
    (1) For an unflared bottom, the pressure at the chine is 0.75 times 
the pressure at the keel, and the pressures between the keel and chine 
vary linearly, in accordance with figure 3 of appendix I of this part. 
The pressure at the keel (p.s.i.) is computed as follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.007

where--

Pk=pressure (p.s.i.) at the keel;
C2=0.00213;
K2=hull station weighing factor, in accordance with figure 2 
          of appendix I of this part;
VS1=seaplane stalling speed (knots) at the design water 
          takeoff weight with flaps extended in the appropriate takeoff 
          position; and
[beta]K=angle of dead rise at keel, in accordance with figure 
          1 of appendix I of this part.

    (2) For a flared bottom, the pressure at the beginning of the flare 
is the same as that for an unflared bottom, and the pressure between the 
chine and the beginning of the flare varies linearly, in accordance with 
figure 3 of appendix I of this part. The pressure distribution is the 
same as that prescribed in paragraph (b)(1) of this section for an 
unflared bottom except that the pressure at the chine is computed as 
follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.008

where--

Pch=pressure (p.s.i.) at the chine;
C3=0.0016;
K2=hull station weighing factor, in accordance with figure 2 
          of appendix I of this part;
VS1=seaplane stalling speed (knots) at the design water 
          takeoff weight with flaps extended in the appropriate takeoff 
          position; and
[beta]=angle of dead rise at appropriate station.

    The area over which these pressures are applied must simulate 
pressures occurring during high localized impacts on the hull or float, 
but need not extend over an area that would induce critical stresses in 
the frames or in the overall structure.
    (c) Distributed pressures. For the design of the frames, keel, and 
chine structure, the following pressure distributions apply:
    (1) Symmetrical pressures are computed as follows:
    [GRAPHIC] [TIFF OMITTED] TC28SE91.009
    
where--

P=pressure (p.s.i.);
C4=0.078 C1 (with C1 computed under 
          Sec. 23.527);
K2=hull station weighing factor, determined in accordance 
          with figure 2 of appendix I of this part;

[[Page 230]]

VS0=seaplane stalling speed (knots) with landing flaps 
          extended in the appropriate position and with no slipstream 
          effect; and
[beta]=angle of dead rise at appropriate station.

    (2) The unsymmetrical pressure distribution consists of the 
pressures prescribed in paragraph (c)(1) of this section on one side of 
the hull or main float centerline and one-half of that pressure on the 
other side of the hull or main float centerline, in accordance with 
figure 3 of appendix I of this part.
    (3) These pressures are uniform and must be applied simultaneously 
over the entire hull or main float bottom. The loads obtained must be 
carried into the sidewall structure of the hull proper, but need not be 
transmitted in a fore and aft direction as shear and bending loads.

[Doc. No. 26269, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]



Sec. 23.535  Auxiliary float loads.

    (a) General. Auxiliary floats and their attachments and supporting 
structures must be designed for the conditions prescribed in this 
section. In the cases specified in paragraphs (b) through (e) of this 
section, the prescribed water loads may be distributed over the float 
bottom to avoid excessive local loads, using bottom pressures not less 
than those prescribed in paragraph (g) of this section.
    (b) Step loading. The resultant water load must be applied in the 
plane of symmetry of the float at a point three-fourths of the distance 
from the bow to the step and must be perpendicular to the keel. The 
resultant limit load is computed as follows, except that the value of L 
need not exceed three times the weight of the displaced water when the 
float is completely submerged:
[GRAPHIC] [TIFF OMITTED] TC28SE91.010

where--

L=limit load (lbs.);
C5=0.0053;
VS0=seaplane stalling speed (knots) with landing flaps 
          extended in the appropriate position and with no slipstream 
          effect;
W=seaplane design landing weight in pounds;
[beta]s=angle of dead rise at a station \3/4\ of the distance from the 
          bow to the step, but need not be less than 15 degrees; and
ry=ratio of the lateral distance between the center of 
          gravity and the plane of symmetry of the float to the radius 
          of gyration in roll.

    (c) Bow loading. The resultant limit load must be applied in the 
plane of symmetry of the float at a point one-fourth of the distance 
from the bow to the step and must be perpendicular to the tangent to the 
keel line at that point. The magnitude of the resultant load is that 
specified in paragraph (b) of this section.
    (d) Unsymmetrical step loading. The resultant water load consists of 
a component equal to 0.75 times the load specified in paragraph (a) of 
this section and a side component equal to 0.025 tan [beta] times the 
load specified in paragraph (b) of this section. The side load must be 
applied perpendicularly to the plane of symmetry of the float at a point 
midway between the keel and the chine.
    (e) Unsymmetrical bow loading. The resultant water load consists of 
a component equal to 0.75 times the load specified in paragraph (b) of 
this section and a side component equal to 0.25 tan [beta] times the 
load specified in paragraph (c) of this section. The side load must be 
applied perpendicularly to the plane of symmetry at a point midway 
between the keel and the chine.
    (f) Immersed float condition. The resultant load must be applied at 
the centroid of the cross section of the float at a point one-third of 
the distance from the bow to the step. The limit load components are as 
follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.011

where--

P=mass density of water (slugs/ft.\3\)
V=volume of float (ft.\3\);
CX=coefficient of drag force, equal to 0.133;
Cy=coefficient of side force, equal to 0.106;

[[Page 231]]

K=0.8, except that lower values may be used if it is shown that the 
          floats are incapable of submerging at a speed of 0.8 
          Vso in normal operations;
Vso=seaplane stalling speed (knots) with landing flaps 
          extended in the appropriate position and with no slipstream 
          effect; and
g=acceleration due to gravity (ft/sec\2\).

    (g) Float bottom pressures. The float bottom pressures must be 
established under Sec. 23.533, except that the value of K2 
in the formulae may be taken as 1.0. The angle of dead rise to be used 
in determining the float bottom pressures is set forth in paragraph (b) 
of this section.

[Doc. No. 26269, 58 FR 42162, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]



Sec. 23.537  Seawing loads.

    Seawing design loads must be based on applicable test data.

[Doc. No. 26269, 58 FR 42163, Aug. 6, 1993]

                      Emergency Landing Conditions



Sec. 23.561  General.

    (a) The airplane, although it may be damaged in emergency landing 
conditions, must be designed as prescribed in this section to protect 
each occupant under those conditions.
    (b) The structure must be designed to give each occupant every 
reasonable chance of escaping serious injury when--
    (1) Proper use is made of the seats, safety belts, and shoulder 
harnesses provided for in the design;
    (2) The occupant experiences the static inertia loads corresponding 
to the following ultimate load factors--
    (i) Upward, 3.0g for normal, utility, and commuter category 
airplanes, or 4.5g for acrobatic category airplanes;
    (ii) Forward, 9.0g;
    (iii) Sideward, 1.5g; and
    (iv) Downward, 6.0g when certification to the emergency exit 
provisions of Sec. 23.807(d)(4) is requested; and
    (3) The items of mass within the cabin, that could injure an 
occupant, experience the static inertia loads corresponding to the 
following ultimate load factors--
    (i) Upward, 3.0g;
    (ii) Forward, 18.0g; and
    (iii) Sideward, 4.5g.
    (c) Each airplane with retractable landing gear must be designed to 
protect each occupant in a landing--
    (1) With the wheels retracted;
    (2) With moderate descent velocity; and
    (3) Assuming, in the absence of a more rational analysis--
    (i) A downward ultimate inertia force of 3 g; and
    (ii) A coefficient of friction of 0.5 at the ground.
    (d) If it is not established that a turnover is unlikely during an 
emergency landing, the structure must be designed to protect the 
occupants in a complete turnover as follows:
    (1) The likelihood of a turnover may be shown by an analysis 
assuming the following conditions--
    (i) The most adverse combination of weight and center of gravity 
position;
    (ii) Longitudinal load factor of 9.0g;
    (iii) Vertical load factor of 1.0g; and
    (iv) For airplanes with tricycle landing gear, the nose wheel strut 
failed with the nose contacting the ground.
    (2) For determining the loads to be applied to the inverted airplane 
after a turnover, an upward ultimate inertia load factor of 3.0g and a 
coefficient of friction with the ground of 0.5 must be used.
    (e) Except as provided in Sec. 23.787(c), the supporting structure 
must be designed to restrain, under loads up to those specified in 
paragraph (b)(3) of this section, each item of mass that could injure an 
occupant if it came loose in a minor crash landing.
    (1) For engines mounted inside the fuselage, aft of the cabin, it 
must be shown by test or analysis that the engine and attached 
accessories, and the engine mounting structure--
    (i) Can withstand a forward acting static ultimate inertia load 
factor of 18.0 g plus the maximum takeoff engine thrust; or
    (ii) The airplane structure is designed to preclude the engine and 
its attached accessories from entering or protruding into the cabin 
should the engine mounts fail.

[[Page 232]]

    (2) [Reserved]

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13090, Aug. 13, 1969; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987; Amdt. 
23-36, 53 FR 30812, Aug. 15, 1988; Amdt. 23-46, 59 FR 25772, May 17, 
1994; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996; Amdt. 23-62, 76 FR 75756, 
Dec. 2, 2011]



Sec. 23.562  Emergency landing dynamic conditions.

    (a) Each seat/restraint system for use in a normal, utility, or 
acrobatic category airplane, or in a commuter category jet airplane, 
must be designed to protect each occupant during an emergency landing 
when--
    (1) Proper use is made of seats, safety belts, and shoulder 
harnesses provided for in the design; and
    (2) The occupant is exposed to the loads resulting from the 
conditions prescribed in this section.
    (b) Except for those seat/restraint systems that are required to 
meet paragraph (d) of this section, each seat/restraint system for crew 
or passenger occupancy in a normal, utility, or acrobatic category 
airplane, or in a commuter category jet airplane, must successfully 
complete dynamic tests or be demonstrated by rational analysis supported 
by dynamic tests, in accordance with each of the following conditions. 
These tests must be conducted with an occupant simulated by an 
anthropomorphic test dummy (ATD) defined by 49 CFR part 572, subpart B, 
or an FAA-approved equivalent, with a nominal weight of 170 pounds and 
seated in the normal upright position.
    (1) For the first test, the change in velocity may not be less than 
31 feet per second. The seat/restraint system must be oriented in its 
nominal position with respect to the airplane and with the horizontal 
plane of the airplane pitched up 60 degrees, with no yaw, relative to 
the impact vector. For seat/restraint systems to be installed in the 
first row of the airplane, peak deceleration must occur in not more than 
0.05 seconds after impact and must reach a minimum of 19g. For all other 
seat/restraint systems, peak deceleration must occur in not more than 
0.06 seconds after impact and must reach a minimum of 15g.
    (2) For the second test, the change in velocity may not be less than 
42 feet per second. The seat/restraint system must be oriented in its 
nominal position with respect to the airplane and with the vertical 
plane of the airplane yawed 10 degrees, with no pitch, relative to the 
impact vector in a direction that results in the greatest load on the 
shoulder harness. For seat/restraint systems to be installed in the 
first row of the airplane, peak deceleration must occur in not more than 
0.05 seconds after impact and must reach a minimum of 26g. For all other 
seat/restraint systems, peak deceleration must occur in not more than 
0.06 seconds after impact and must reach a minimum of 21g.
    (3) To account for floor warpage, the floor rails or attachment 
devices used to attach the seat/restraint system to the airframe 
structure must be preloaded to misalign with respect to each other by at 
least 10 degrees vertically (i.e., pitch out of parallel) and one of the 
rails or attachment devices must be preloaded to misalign by 10 degrees 
in roll prior to conducting the test defined by paragraph (b)(2) of this 
section.
    (c) Compliance with the following requirements must be shown during 
the dynamic tests conducted in accordance with paragraph (b) of this 
section:
    (1) The seat/restraint system must restrain the ATD although seat/
restraint system components may experience deformation, elongation, 
displacement, or crushing intended as part of the design.
    (2) The attachment between the seat/restraint system and the test 
fixture must remain intact, although the seat structure may have 
deformed.
    (3) Each shoulder harness strap must remain on the ATD's shoulder 
during the impact.
    (4) The safety belt must remain on the ATD's pelvis during the 
impact.
    (5) The results of the dynamic tests must show that the occupant is 
protected from serious head injury.
    (i) When contact with adjacent seats, structure, or other items in 
the cabin can occur, protection must be provided so that the head impact 
does not exceed a head injury criteria (HIC) of 1,000.

[[Page 233]]

    (ii) The value of HIC is defined as--
    [GRAPHIC] [TIFF OMITTED] TR02DE11.083
    
Where--

t1 is the initial integration time, expressed in seconds, 
          t2 is the final integration time, expressed in 
          seconds, and a(t) is the total acceleration vs. time curve for 
          the head strike expressed as a multiple of g (units of 
          gravity).

    (iii) Compliance with the HIC limit must be demonstrated by 
measuring the head impact during dynamic testing as prescribed in 
paragraphs (b)(1) and (b)(2) of this section or by a separate showing of 
compliance with the head injury criteria using test or analysis 
procedures.
    (6) Loads in individual shoulder harness straps may not exceed 1,750 
pounds. If dual straps are used for retaining the upper torso, the total 
strap loads may not exceed 2,000 pounds.
    (7) The compression load measured between the pelvis and the lumbar 
spine of the ATD may not exceed 1,500 pounds.
    (d) For all single-engine airplanes with a VSO of more 
than 61 knots at maximum weight, and those multiengine airplanes of 
6,000 pounds or less maximum weight with a VSO of more than 
61 knots at maximum weight that do not comply with Sec. 23.67(a)(1);
    (1) The ultimate load factors of Sec. 23.561(b) must be increased 
by multiplying the load factors by the square of the ratio of the 
increased stall speed to 61 knots. The increased ultimate load factors 
need not exceed the values reached at a VS0 of 79 knots. The 
upward ultimate load factor for acrobatic category airplanes need not 
exceed 5.0g.
    (2) The seat/restraint system test required by paragraph (b)(1) of 
this section must be conducted in accordance with the following 
criteria:
    (i) The change in velocity may not be less than 31 feet per second.
    (ii)(A) The peak deceleration (gp) of 19g and 15g must be 
increased and multiplied by the square of the ratio of the increased 
stall speed to 61 knots:

gp=19.0 (VS0/61)\2\ or gp=15.0 
(VS0/61)\2\

    (B) The peak deceleration need not exceed the value reached at a 
VS0 of 79 knots.
    (iii) The peak deceleration must occur in not more than time 
(tr), which must be computed as follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.013

where--

gp=The peak deceleration calculated in accordance with 
          paragraph (d)(2)(ii) of this section
tr=The rise time (in seconds) to the peak deceleration.

    (e) An alternate approach that achieves an equivalent, or greater, 
level of occupant protection to that required by this section may be 
used if substantiated on a rational basis.

[Amdt. 23-36, 53 FR 30812, Aug. 15, 1988, as amended by Amdt. 23-44, 58 
FR 38639, July 19, 1993; Amdt. 23-50, 61 FR 5192, Feb. 9, 1996; Amdt. 
23-62, 76 FR 75756, Dec. 2, 2011]

                           Fatigue Evaluation



Sec. 23.571  Metallic pressurized cabin structures.

    For normal, utility, and acrobatic category airplanes, the strength, 
detail design, and fabrication of the metallic structure of the pressure 
cabin must be evaluated under one of the following:
    (a) A fatigue strength investigation in which the structure is shown 
by tests, or by analysis supported by test evidence, to be able to 
withstand the repeated loads of variable magnitude expected in service; 
or
    (b) A fail safe strength investigation, in which it is shown by 
analysis, tests, or both that catastrophic failure of the structure is 
not probable after fatigue

[[Page 234]]

failure, or obvious partial failure, of a principal structural element, 
and that the remaining structures are able to withstand a static 
ultimate load factor of 75 percent of the limit load factor at VC, 
considering the combined effects of normal operating pressures, expected 
external aerodynamic pressures, and flight loads. These loads must be 
multiplied by a factor of 1.15 unless the dynamic effects of failure 
under static load are otherwise considered.
    (c) The damage tolerance evaluation of Sec. 23.573(b).
    (d) If certification for operation above 41,000 feet is requested, a 
damage tolerance evaluation of the fuselage pressure boundary per Sec. 
23.573(b) must be conducted.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31821, Nov. 19, 1973; Amdt. 23-45, 58 FR 42163, Aug. 6, 1993; 
Amdt. 23-48, 61 FR 5147, Feb. 9, 1996; Amdt. 23-62, 76 FR 75756, Dec. 2, 
2011]



Sec. 23.572  Metallic wing, empennage, and associated structures.

    (a) For normal, utility, and acrobatic category airplanes, the 
strength, detail design, and fabrication of those parts of the airframe 
structure whose failure would be catastrophic must be evaluated under 
one of the following unless it is shown that the structure, operating 
stress level, materials and expected uses are comparable, from a fatigue 
standpoint, to a similar design that has had extensive satisfactory 
service experience:
    (1) A fatigue strength investigation in which the structure is shown 
by tests, or by analysis supported by test evidence, to be able to 
withstand the repeated loads of variable magnitude expected in service; 
or
    (2) A fail-safe strength investigation in which it is shown by 
analysis, tests, or both, that catastrophic failure of the structure is 
not probable after fatigue failure, or obvious partial failure, of a 
principal structural element, and that the remaining structure is able 
to withstand a static ultimate load factor of 75 percent of the critical 
limit load factor at Vc. These loads must be multiplied by a factor of 
1.15 unless the dynamic effects of failure under static load are 
otherwise considered.
    (3) The damage tolerance evaluation of Sec. 23.573(b).
    (b) Each evaluation required by this section must--
    (1) Include typical loading spectra (e.g. taxi, ground-air-ground 
cycles, maneuver, gust);
    (2) Account for any significant effects due to the mutual influence 
of aerodynamic surfaces; and
    (3) Consider any significant effects from propeller slipstream 
loading, and buffet from vortex impingements.

[Amdt. 23-7, 34 FR 13090, Aug. 13, 1969, as amended by Amdt. 23-14, 38 
FR 31821, Nov. 19, 1973; Amdt. 23-34, 52 FR 1830, Jan. 15, 1987; Amdt. 
23-38, 54 FR 39511, Sept. 26, 1989; Amdt. 23-45, 58 FR 42163, Aug. 6, 
1993; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996]



Sec. 23.573  Damage tolerance and fatigue evaluation of structure.

    (a) Composite airframe structure. Composite airframe structure must 
be evaluated under this paragraph instead of Sec. Sec. 23.571 and 
23.572. The applicant must evaluate the composite airframe structure, 
the failure of which would result in catastrophic loss of the airplane, 
in each wing (including canards, tandem wings, and winglets), empennage, 
their carrythrough and attaching structure, moveable control surfaces 
and their attaching structure fuselage, and pressure cabin using the 
damage-tolerance criteria prescribed in paragraphs (a)(1) through (a)(4) 
of this section unless shown to be impractical. If the applicant 
establishes that damage-tolerance criteria is impractical for a 
particular structure, the structure must be evaluated in accordance with 
paragraphs (a)(1) and (a)(6) of this section. Where bonded joints are 
used, the structure must also be evaluated in accordance with paragraph 
(a)(5) of this section. The effects of material variability and 
environmental conditions on the strength and durability properties of 
the composite materials must be accounted for in the evaluations 
required by this section.
    (1) It must be demonstrated by tests, or by analysis supported by 
tests, that the structure is capable of carrying ultimate load with 
damage up to the threshold of detectability considering the inspection 
procedures employed.

[[Page 235]]

    (2) The growth rate or no-growth of damage that may occur from 
fatigue, corrosion, manufacturing flaws or impact damage, under repeated 
loads expected in service, must be established by tests or analysis 
supported by tests.
    (3) The structure must be shown by residual strength tests, or 
analysis supported by residual strength tests, to be able to withstand 
critical limit flight loads, considered as ultimate loads, with the 
extent of detectable damage consistent with the results of the damage 
tolerance evaluations. For pressurized cabins, the following loads must 
be withstood:
    (i) Critical limit flight loads with the combined effects of normal 
operating pressure and expected external aerodynamic pressures.
    (ii) The expected external aerodynamic pressures in 1g flight 
combined with a cabin differential pressure equal to 1.1 times the 
normal operating differential pressure without any other load.
    (4) The damage growth, between initial detectability and the value 
selected for residual strength demonstrations, factored to obtain 
inspection intervals, must allow development of an inspection program 
suitable for application by operation and maintenance personnel.
    (5) For any bonded joint, the failure of which would result in 
catastrophic loss of the airplane, the limit load capacity must be 
substantiated by one of the following methods--
    (i) The maximum disbonds of each bonded joint consistent with the 
capability to withstand the loads in paragraph (a)(3) of this section 
must be determined by analysis, tests, or both. Disbonds of each bonded 
joint greater than this must be prevented by design features; or
    (ii) Proof testing must be conducted on each production article that 
will apply the critical limit design load to each critical bonded joint; 
or
    (iii) Repeatable and reliable non-destructive inspection techniques 
must be established that ensure the strength of each joint.
    (6) Structural components for which the damage tolerance method is 
shown to be impractical must be shown by component fatigue tests, or 
analysis supported by tests, to be able to withstand the repeated loads 
of variable magnitude expected in service. Sufficient component, 
subcomponent, element, or coupon tests must be done to establish the 
fatigue scatter factor and the environmental effects. Damage up to the 
threshold of detectability and ultimate load residual strength 
capability must be considered in the demonstration.
    (b) Metallic airframe structure. If the applicant elects to use 
Sec. 23.571(c) or Sec. 23.572(a)(3), then the damage tolerance 
evaluation must include a determination of the probable locations and 
modes of damage due to fatigue, corrosion, or accidental damage. Damage 
at multiple sites due to fatigue must be included where the design is 
such that this type of damage can be expected to occur. The evaluation 
must incorporate repeated load and static analyses supported by test 
evidence. The extent of damage for residual strength evaluation at any 
time within the operational life of the airplane must be consistent with 
the initial detectability and subsequent growth under repeated loads. 
The residual strength evaluation must show that the remaining structure 
is able to withstand critical limit flight loads, considered as 
ultimate, with the extent of detectable damage consistent with the 
results of the damage tolerance evaluations. For pressurized cabins, the 
following load must be withstood:
    (1) The normal operating differential pressure combined with the 
expected external aerodynamic pressures applied simultaneously with the 
flight loading conditions specified in this part, and
    (2) The expected external aerodynamic pressures in 1g flight 
combined with a cabin differential pressure equal to 1.1 times the 
normal operating differential pressure without any other load.

[Doc. No. 26269, 58 FR 42163, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993, 
as amended by Amdt. 23-48, 61 FR 5147, Feb. 9, 1996; 73 FR 19746, Apr. 
11, 2008]



Sec. 23.574  Metallic damage tolerance and fatigue evaluation of
commuter category airplanes.

    For commuter category airplanes--

[[Page 236]]

    (a) Metallic damage tolerance. An evaluation of the strength, detail 
design, and fabrication must show that catastrophic failure due to 
fatigue, corrosion, defects, or damage will be avoided throughout the 
operational life of the airplane. This evaluation must be conducted in 
accordance with the provisions of Sec. 23.573, except as specified in 
paragraph (b) of this section, for each part of the structure that could 
contribute to a catastrophic failure.
    (b) Fatigue (safe-life) evaluation. Compliance with the damage 
tolerance requirements of paragraph (a) of this section is not required 
if the applicant establishes that the application of those requirements 
is impractical for a particular structure. This structure must be shown, 
by analysis supported by test evidence, to be able to withstand the 
repeated loads of variable magnitude expected during its service life 
without detectable cracks. Appropriate safe-life scatter factors must be 
applied.

[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]



Sec. 23.575  Inspections and other procedures.

    Each inspection or other procedure, based on an evaluation required 
by Sec. Sec. 23.571, 23.572, 23.573 or 23.574, must be established to 
prevent catastrophic failure and must be included in the Limitations 
Section of the Instructions for Continued Airworthiness required by 
Sec. 23.1529.

[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]



                    Subpart D_Design and Construction



Sec. 23.601  General.

    The suitability of each questionable design detail and part having 
an important bearing on safety in operations, must be established by 
tests.



Sec. 23.603  Materials and workmanship.

    (a) The suitability and durability of materials used for parts, the 
failure of which could adversely affect safety, must--
    (1) Be established by experience or tests;
    (2) Meet approved specifications that ensure their having the 
strength and other properties assumed in the design data; and
    (3) Take into account the effects of environmental conditions, such 
as temperature and humidity, expected in service.
    (b) Workmanship must be of a high standard.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 
41 FR 55464, Dec. 20, 1976; Amdt. 23-23, 43 FR 50592, Oct. 10, 1978]



Sec. 23.605  Fabrication methods.

    (a) The methods of fabrication used must produce consistently sound 
structures. If a fabrication process (such as gluing, spot welding, or 
heat-treating) requires close control to reach this objective, the 
process must be performed under an approved process specification.
    (b) Each new aircraft fabrication method must be substantiated by a 
test program.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-23, 43 FR 50592, Oct. 10, 1978]



Sec. 23.607  Fasteners.

    (a) Each removable fastener must incorporate two retaining devices 
if the loss of such fastener would preclude continued safe flight and 
landing.
    (b) Fasteners and their locking devices must not be adversely 
affected by the environmental conditions associated with the particular 
installation.
    (c) No self-locking nut may be used on any bolt subject to rotation 
in operation unless a non-friction locking device is used in addition to 
the self-locking device.

[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]



Sec. 23.609  Protection of structure.

    Each part of the structure must--
    (a) Be suitably protected against deterioration or loss of strength 
in service due to any cause, including--
    (1) Weathering;
    (2) Corrosion; and
    (3) Abrasion; and
    (b) Have adequate provisions for ventilation and drainage.



Sec. 23.611  Accessibility provisions.

    For each part that requires maintenance, inspection, or other 
servicing,

[[Page 237]]

appropriate means must be incorporated into the aircraft design to allow 
such servicing to be accomplished.

[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]



Sec. 23.613  Material strength properties and design values.

    (a) Material strength properties must be based on enough tests of 
material meeting specifications to establish design values on a 
statistical basis.
    (b) Design values must be chosen to minimize the probability of 
structural failure due to material variability. Except as provided in 
paragraph (e) of this section, compliance with this paragraph must be 
shown by selecting design values that ensure material strength with the 
following probability:
    (1) Where applied loads are eventually distributed through a single 
member within an assembly, the failure of which would result in loss of 
structural integrity of the component; 99 percent probability with 95 
percent confidence.
    (2) For redundant structure, in which the failure of individual 
elements would result in applied loads being safely distributed to other 
load carrying members; 90 percent probability with 95 percent 
confidence.
    (c) The effects of temperature on allowable stresses used for design 
in an essential component or structure must be considered where thermal 
effects are significant under normal operating conditions.
    (d) The design of the structure must minimize the probability of 
catastrophic fatigue failure, particularly at points of stress 
concentration.
    (e) Design values greater than the guaranteed minimums required by 
this section may be used where only guaranteed minimum values are 
normally allowed if a ``premium selection'' of the material is made in 
which a specimen of each individual item is tested before use to 
determine that the actual strength properties of that particular item 
will equal or exceed those used in design.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-23, 43 FR 50592, Oct. 30, 1978; Amdt. 23-45, 58 FR 
42163, Aug. 6, 1993]



Sec. 23.619  Special factors.

    The factor of safety prescribed in Sec. 23.303 must be multiplied 
by the highest pertinent special factors of safety prescribed in 
Sec. Sec. 23.621 through 23.625 for each part of the structure whose 
strength is--
    (a) Uncertain;
    (b) Likely to deteriorate in service before normal replacement; or
    (c) Subject to appreciable variability because of uncertainties in 
manufacturing processes or inspection methods.

[Amdt. 23-7, 34 FR 13091, Aug. 13, 1969]



Sec. 23.621  Casting factors.

    (a) General. The factors, tests, and inspections specified in 
paragraphs (b) through (d) of this section must be applied in addition 
to those necessary to establish foundry quality control. The inspections 
must meet approved specifications. Paragraphs (c) and (d) of this 
section apply to any structural castings except castings that are 
pressure tested as parts of hydraulic or other fluid systems and do not 
support structural loads.
    (b) Bearing stresses and surfaces. The casting factors specified in 
paragraphs (c) and (d) of this section--
    (1) Need not exceed 1.25 with respect to bearing stresses regardless 
of the method of inspection used; and
    (2) Need not be used with respect to the bearing surfaces of a part 
whose bearing factor is larger than the applicable casting factor.
    (c) Critical castings. For each casting whose failure would preclude 
continued safe flight and landing of the airplane or result in serious 
injury to occupants, the following apply:
    (1) Each critical casting must either--
    (i) Have a casting factor of not less than 1.25 and receive 100 
percent inspection by visual, radiographic, and either magnetic 
particle, penetrant or other approved equivalent non-destructive 
inspection method; or
    (ii) Have a casting factor of not less than 2.0 and receive 100 
percent visual inspection and 100 percent approved non-destructive 
inspection. When an approved quality control procedure is established 
and an acceptable statistical analysis supports reduction, non-

[[Page 238]]

destructive inspection may be reduced from 100 percent, and applied on a 
sampling basis.
    (2) For each critical casting with a casting factor less than 1.50, 
three sample castings must be static tested and shown to meet--
    (i) The strength requirements of Sec. 23.305 at an ultimate load 
corresponding to a casting factor of 1.25; and
    (ii) The deformation requirements of Sec. 23.305 at a load of 1.15 
times the limit load.
    (3) Examples of these castings are structural attachment fittings, 
parts of flight control systems, control surface hinges and balance 
weight attachments, seat, berth, safety belt, and fuel and oil tank 
supports and attachments, and cabin pressure valves.
    (d) Non-critical castings. For each casting other than those 
specified in paragraph (c) or (e) of this section, the following apply:
    (1) Except as provided in paragraphs (d)(2) and (3) of this section, 
the casting factors and corresponding inspections must meet the 
following table:

------------------------------------------------------------------------
              Casting factor                         Inspection
------------------------------------------------------------------------
2.0 or more..............................  100 percent visual.
Less than 2.0 but more than 1.5..........  100 percent visual, and
                                            magnetic particle or
                                            penetrant or equivalent
                                            nondestructive inspection
                                            methods.
1.25 through 1.50........................  100 percent visual, magnetic
                                            particle or penetrant, and
                                            radiographic, or approved
                                            equivalent nondestructive
                                            inspection methods.
------------------------------------------------------------------------

    (2) The percentage of castings inspected by nonvisual methods may be 
reduced below that specified in subparagraph (d)(1) of this section when 
an approved quality control procedure is established.
    (3) For castings procured to a specification that guarantees the 
mechanical properties of the material in the casting and provides for 
demonstration of these properties by test of coupons cut from the 
castings on a sampling basis--
    (i) A casting factor of 1.0 may be used; and
    (ii) The castings must be inspected as provided in paragraph (d)(1) 
of this section for casting factors of ``1.25 through 1.50'' and tested 
under paragraph (c)(2) of this section.
    (e) Non-structural castings. Castings used for non-structural 
purposes do not require evaluation, testing or close inspection.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42164, Aug. 6, 1993]



Sec. 23.623  Bearing factors.

    (a) Each part that has clearance (free fit), and that is subject to 
pounding or vibration, must have a bearing factor large enough to 
provide for the effects of normal relative motion.
    (b) For control surface hinges and control system joints, compliance 
with the factors prescribed in Sec. Sec. 23.657 and 23.693, 
respectively, meets paragraph (a) of this section.

[Amdt. 23-7, 34 FR 13091, Aug. 13, 1969]



Sec. 23.625  Fitting factors.

    For each fitting (a part or terminal used to join one structural 
member to another), the following apply:
    (a) For each fitting whose strength is not proven by limit and 
ultimate load tests in which actual stress conditions are simulated in 
the fitting and surrounding structures, a fitting factor of at least 
1.15 must be applied to each part of--
    (1) The fitting;
    (2) The means of attachment; and
    (3) The bearing on the joined members.
    (b) No fitting factor need be used for joint designs based on 
comprehensive test data (such as continuous joints in metal plating, 
welded joints, and scarf joints in wood).
    (c) For each integral fitting, the part must be treated as a fitting 
up to the point at which the section properties become typical of the 
member.
    (d) For each seat, berth, safety belt, and harness, its attachment 
to the structure must be shown, by analysis, tests, or both, to be able 
to withstand the inertia forces prescribed in Sec. 23.561 multiplied by 
a fitting factor of 1.33.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13091, Aug. 13, 1969]

[[Page 239]]



Sec. 23.627  Fatigue strength.

    The structure must be designed, as far as practicable, to avoid 
points of stress concentration where variable stresses above the fatigue 
limit are likely to occur in normal service.



Sec. 23.629  Flutter.

    (a) It must be shown by the methods of paragraph (b) and either 
paragraph (c) or (d) of this section, that the airplane is free from 
flutter, control reversal, and divergence for any condition of operation 
within the limit V-n envelope and at all speeds up to the speed 
specified for the selected method. In addition--
    (1) Adequate tolerances must be established for quantities which 
affect flutter, including speed, damping, mass balance, and control 
system stiffness; and
    (2) The natural frequencies of main structural components must be 
determined by vibration tests or other approved methods.
    (b) Flight flutter tests must be made to show that the airplane is 
free from flutter, control reversal and divergence and to show that--
    (1) Proper and adequate attempts to induce flutter have been made 
within the speed range up to VD/MD, or 
VDF/MDF for jets;
    (2) The vibratory response of the structure during the test 
indicates freedom from flutter;
    (3) A proper margin of damping exists at VD/
MD, or VDF/MDF for jets; and
    (4) As VD/MD (or VDF/MDF 
for jets) is approached, there is no large or rapid reduction in 
damping.
    (c) Any rational analysis used to predict freedom from flutter, 
control reversal and divergence must cover all speeds up to 1.2 
VD/1.2 MD, limited to Mach 1.0 for subsonic 
airplanes.
    (d) Compliance with the rigidity and mass balance criteria (pages 4-
12), in Airframe and Equipment Engineering Report No. 45 (as corrected) 
``Simplified Flutter Prevention Criteria'' (published by the Federal 
Aviation Administration) may be accomplished to show that the airplane 
is free from flutter, control reversal, or divergence if--
    (1) VD/MD for the airplane is less than 260 
knots (EAS) and less than Mach 0.5,
    (2) The wing and aileron flutter prevention criteria, as represented 
by the wing torsional stiffness and aileron balance criteria, are 
limited in use to airplanes without large mass concentrations (such as 
engines, floats, or fuel tanks in outer wing panels) along the wing 
span, and
    (3) The airplane--
    (i) Does not have a T-tail or other unconventional tail 
configurations;
    (ii) Does not have unusual mass distributions or other 
unconventional design features that affect the applicability of the 
criteria, and
    (iii) Has fixed-fin and fixed-stabilizer surfaces.
    (e) For turbopropeller-powered airplanes, the dynamic evaluation 
must include--
    (1) Whirl mode degree of freedom which takes into account the 
stability of the plane of rotation of the propeller and significant 
elastic, inertial, and aerodynamic forces, and
    (2) Propeller, engine, engine mount, and airplane structure 
stiffness and damping variations appropriate to the particular 
configuration.
    (f) Freedom from flutter, control reversal, and divergence up to 
VD/MD must be shown as follows:
    (1) For airplanes that meet the criteria of paragraphs (d)(1) 
through (d)(3) of this section, after the failure, malfunction, or 
disconnection of any single element in any tab control system.
    (2) For airplanes other than those described in paragraph (f)(1) of 
this section, after the failure, malfunction, or disconnection of any 
single element in the primary flight control system, any tab control 
system, or any flutter damper.
    (g) For airplanes showing compliance with the fail-safe criteria of 
Sec. Sec. 23.571 and 23.572, the airplane must be shown by analysis to 
be free from flutter up to VD/MD after fatigue 
failure, or obvious partial failure, of a principal structural element.
    (h) For airplanes showing compliance with the damage tolerance 
criteria of Sec. 23.573, the airplane must be shown by analysis to be 
free from flutter up to VD/MD with the extent of 
damage for which residual strength is demonstrated.

[[Page 240]]

    (i) For modifications to the type design that could affect the 
flutter characteristics, compliance with paragraph (a) of this section 
must be shown, except that analysis based on previously approved data 
may be used alone to show freedom from flutter, control reversal and 
divergence, for all speeds up to the speed specified for the selected 
method.

[Amdt. 23-23, 43 FR 50592, Oct. 30, 1978, as amended by Amdt. 23-31, 49 
FR 46867, Nov. 28, 1984; Amdt. 23-45, 58 FR 42164, Aug. 6, 1993; 58 FR 
51970, Oct. 5, 1993; Amdt. 23-48, 61 FR 5148, Feb. 9, 1996; Amdt. 23-62, 
76 FR 75756, Dec. 2, 2011]

                                  Wings



Sec. 23.641  Proof of strength.

    The strength of stressed-skin wings must be proven by load tests or 
by combined structural analysis and load tests.

                            Control Surfaces



Sec. 23.651  Proof of strength.

    (a) Limit load tests of control surfaces are required. These tests 
must include the horn or fitting to which the control system is 
attached.
    (b) In structural analyses, rigging loads due to wire bracing must 
be accounted for in a rational or conservative manner.



Sec. 23.655  Installation.

    (a) Movable surfaces must be installed so that there is no 
interference between any surfaces, their bracing, or adjacent fixed 
structure, when one surface is held in its most critical clearance 
positions and the others are operated through their full movement.
    (b) If an adjustable stabilizer is used, it must have stops that 
will limit its range of travel to that allowing safe flight and landing.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42164, Aug. 6, 1993]



Sec. 23.657  Hinges.

    (a) Control surface hinges, except ball and roller bearing hinges, 
must have a factor of safety of not less than 6.67 with respect to the 
ultimate bearing strength of the softest material used as a bearing.
    (b) For ball or roller bearing hinges, the approved rating of the 
bearing may not be exceeded.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-48, 
61 FR 5148, Feb. 9, 1996]



Sec. 23.659  Mass balance.

    The supporting structure and the attachment of concentrated mass 
balance weights used on control surfaces must be designed for--
    (a) 24 g normal to the plane of the control surface;
    (b) 12 g fore and aft; and
    (c) 12 g parallel to the hinge line.

                             Control Systems



Sec. 23.671  General.

    (a) Each control must operate easily, smoothly, and positively 
enough to allow proper performance of its functions.
    (b) Controls must be arranged and identified to provide for 
convenience in operation and to prevent the possibility of confusion and 
subsequent inadvertent operation.



Sec. 23.672  Stability augmentation and automatic and power-operated
systems.

    If the functioning of stability augmentation or other automatic or 
power-operated systems is necessary to show compliance with the flight 
characteristics requirements of this part, such systems must comply with 
Sec. 23.671 and the following:
    (a) A warning, which is clearly distinguishable to the pilot under 
expected flight conditions without requiring the pilot's attention, must 
be provided for any failure in the stability augmentation system or in 
any other automatic or power-operated system that could result in an 
unsafe condition if the pilot was not aware of the failure. Warning 
systems must not activate the control system.
    (b) The design of the stability augmentation system or of any other 
automatic or power-operated system must permit initial counteraction of 
failures without requiring exceptional pilot

[[Page 241]]

skill or strength, by either the deactivation of the system or a failed 
portion thereof, or by overriding the failure by movement of the flight 
controls in the normal sense.
    (c) It must be shown that, after any single failure of the stability 
augmentation system or any other automatic or power-operated system--
    (1) The airplane is safely controllable when the failure or 
malfunction occurs at any speed or altitude within the approved 
operating limitations that is critical for the type of failure being 
considered;
    (2) The controllability and maneuverability requirements of this 
part are met within a practical operational flight envelope (for 
example, speed, altitude, normal acceleration, and airplane 
configuration) that is described in the Airplane Flight Manual (AFM); 
and
    (3) The trim, stability, and stall characteristics are not impaired 
below a level needed to permit continued safe flight and landing.

[Doc. No. 26269, 58 FR 42164, Aug. 6, 1993]



Sec. 23.673  Primary flight controls.

    Primary flight controls are those used by the pilot for the 
immediate control of pitch, roll, and yaw.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-48, 
61 FR 5148, Feb. 9, 1996]



Sec. 23.675  Stops.

    (a) Each control system must have stops that positively limit the 
range of motion of each movable aerodynamic surface controlled by the 
system.
    (b) Each stop must be located so that wear, slackness, or takeup 
adjustments will not adversely affect the control characteristics of the 
airplane because of a change in the range of surface travel.
    (c) Each stop must be able to withstand any loads corresponding to 
the design conditions for the control system.

[Amdt. 23-17, 41 FR 55464, Dec. 20, 1976]



Sec. 23.677  Trim systems.

    (a) Proper precautions must be taken to prevent inadvertent, 
improper, or abrupt trim tab operation. There must be means near the 
trim control to indicate to the pilot the direction of trim control 
movement relative to airplane motion. In addition, there must be means 
to indicate to the pilot the position of the trim device with respect to 
both the range of adjustment and, in the case of lateral and directional 
trim, the neutral position. This means must be visible to the pilot and 
must be located and designed to prevent confusion. The pitch trim 
indicator must be clearly marked with a position or range within which 
it has been demonstrated that take-off is safe for all center of gravity 
positions and each flap position approved for takeoff.
    (b) Trimming devices must be designed so that, when any one 
connecting or transmitting element in the primary flight control system 
fails, adequate control for safe flight and landing is available with--
    (1) For single-engine airplanes, the longitudinal trimming devices; 
or
    (2) For multiengine airplanes, the longitudinal and directional 
trimming devices.
    (c) Tab controls must be irreversible unless the tab is properly 
balanced and has no unsafe flutter characteristics. Irreversible tab 
systems must have adequate rigidity and reliability in the portion of 
the system from the tab to the attachment of the irreversible unit to 
the airplane structure.
    (d) It must be demonstrated that the airplane is safely controllable 
and that the pilot can perform all maneuvers and operations necessary to 
effect a safe landing following any probable powered trim system runaway 
that reasonably might be expected in service, allowing for appropriate 
time delay after pilot recognition of the trim system runaway. The 
demonstration must be conducted at critical airplane weights and center 
of gravity positions.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13091, Aug. 13, 1969; Amdt. 23-34, 52 FR 1830, Jan. 15, 1987; Amdt. 
23-42, 56 FR 353, Jan. 3, 1991; Amdt. 23-49, 61 FR 5165, Feb. 9, 1996]



Sec. 23.679  Control system locks.

    If there is a device to lock the control system on the ground or 
water:
    (a) There must be a means to--

[[Page 242]]

    (1) Give unmistakable warning to the pilot when lock is engaged; or
    (2) Automatically disengage the device when the pilot operates the 
primary flight controls in a normal manner.
    (b) The device must be installed to limit the operation of the 
airplane so that, when the device is engaged, the pilot receives 
unmistakable warning at the start of the takeoff.
    (c) The device must have a means to preclude the possibility of it 
becoming inadvertently engaged in flight.

[Doc. No. 26269, 58 FR 42164, Aug. 6, 1993]



Sec. 23.681  Limit load static tests.

    (a) Compliance with the limit load requirements of this part must be 
shown by tests in which--
    (1) The direction of the test loads produces the most severe loading 
in the control system; and
    (2) Each fitting, pulley, and bracket used in attaching the system 
to the main structure is included.
    (b) Compliance must be shown (by analyses or individual load tests) 
with the special factor requirements for control system joints subject 
to angular motion.



Sec. 23.683  Operation tests.

    (a) It must be shown by operation tests that, when the controls are 
operated from the pilot compartment with the system loaded as prescribed 
in paragraph (b) of this section, the system is free from--
    (1) Jamming;
    (2) Excessive friction; and
    (3) Excessive deflection.
    (b) The prescribed test loads are--
    (1) For the entire system, loads corresponding to the limit airloads 
on the appropriate surface, or the limit pilot forces in Sec. 
23.397(b), whichever are less; and
    (2) For secondary controls, loads not less than those corresponding 
to the maximum pilot effort established under Sec. 23.405.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13091, Aug. 13, 1969]



Sec. 23.685  Control system details.

    (a) Each detail of each control system must be designed and 
installed to prevent jamming, chafing, and interference from cargo, 
passengers, loose objects, or the freezing of moisture.
    (b) There must be means in the cockpit to prevent the entry of 
foreign objects into places where they would jam the system.
    (c) There must be means to prevent the slapping of cables or tubes 
against other parts.
    (d) Each element of the flight control system must have design 
features, or must be distinctively and permanently marked, to minimize 
the possibility of incorrect assembly that could result in 
malfunctioning of the control system.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 
41 FR 55464, Dec. 20, 1976]



Sec. 23.687  Spring devices.

    The reliability of any spring device used in the control system must 
be established by tests simulating service conditions unless failure of 
the spring will not cause flutter or unsafe flight characteristics.



Sec. 23.689  Cable systems.

    (a) Each cable, cable fitting, turnbuckle, splice, and pulley used 
must meet approved specifications. In addition--
    (1) No cable smaller than \1/8\ inch diameter may be used in primary 
control systems;
    (2) Each cable system must be designed so that there will be no 
hazardous change in cable tension throughout the range of travel under 
operating conditions and temperature variations; and
    (3) There must be means for visual inspection at each fairlead, 
pulley, terminal, and turnbuckle.
    (b) Each kind and size of pulley must correspond to the cable with 
which it is used. Each pulley must have closely fitted guards to prevent 
the cables from being misplaced or fouled, even when slack. Each pulley 
must lie in the plane passing through the cable so that the cable does 
not rub against the pulley flange.

[[Page 243]]

    (c) Fairleads must be installed so that they do not cause a change 
in cable direction of more than three degrees.
    (d) Clevis pins subject to load or motion and retained only by 
cotter pins may not be used in the control system.
    (e) Turnbuckles must be attached to parts having angular motion in a 
manner that will positively prevent binding throughout the range of 
travel.
    (f) Tab control cables are not part of the primary control system 
and may be less than \1/8\ inch diameter in airplanes that are safely 
controllable with the tabs in the most adverse positions.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13091, Aug. 13, 1969]



Sec. 23.691  Artificial stall barrier system.

    If the function of an artificial stall barrier, for example, stick 
pusher, is used to show compliance with Sec. 23.201(c), the system must 
comply with the following:
    (a) With the system adjusted for operation, the plus and minus 
airspeeds at which downward pitching control will be provided must be 
established.
    (b) Considering the plus and minus airspeed tolerances established 
by paragraph (a) of this section, an airspeed must be selected for the 
activation of the downward pitching control that provides a safe margin 
above any airspeed at which any unsatisfactory stall characteristics 
occur.
    (c) In addition to the stall warning required Sec. 23.07, a warning 
that is clearly distinguishable to the pilot under all expected flight 
conditions without requiring the pilot's attention, must be provided for 
faults that would prevent the system from providing the required 
pitching motion.
    (d) Each system must be designed so that the artificial stall 
barrier can be quickly and positively disengaged by the pilots to 
prevent unwanted downward pitching of the airplane by a quick release 
(emergency) control that meets the requirements of Sec. 23.1329(b).
    (e) A preflight check of the complete system must be established and 
the procedure for this check made available in the Airplane Flight 
Manual (AFM). Preflight checks that are critical to the safety of the 
airplane must be included in the limitations section of the AFM.
    (f) For those airplanes whose design includes an autopilot system:
    (1) A quick release (emergency) control installed in accordance with 
Sec. 23.1329(b) may be used to meet the requirements of paragraph (d), 
of this section, and
    (2) The pitch servo for that system may be used to provide the stall 
downward pitching motion.
    (g) In showing compliance with Sec. 23.1309, the system must be 
evaluated to determine the effect that any announced or unannounced 
failure may have on the continued safe flight and landing of the 
airplane or the ability of the crew to cope with any adverse conditions 
that may result from such failures. This evaluation must consider the 
hazards that would result from the airplane's flight characteristics if 
the system was not provided, and the hazard that may result from 
unwanted downward pitching motion, which could result from a failure at 
airspeeds above the selected stall speed.

[Doc. No. 27806, 61 FR 5165, Feb. 9, 1996]



Sec. 23.693  Joints.

    Control system joints (in push-pull systems) that are subject to 
angular motion, except those in ball and roller bearing systems, must 
have a special factor of safety of not less than 3.33 with respect to 
the ultimate bearing strength of the softest material used as a bearing. 
This factor may be reduced to 2.0 for joints in cable control systems. 
For ball or roller bearings, the approved ratings may not be exceeded.



Sec. 23.697  Wing flap controls.

    (a) Each wing flap control must be designed so that, when the flap 
has been placed in any position upon which compliance with the 
performance requirements of this part is based, the flap will not move 
from that position unless the control is adjusted or is moved by the 
automatic operation of a flap load limiting device.
    (b) The rate of movement of the flaps in response to the operation 
of the pilot's control or automatic device must give satisfactory flight 
and performance characteristics under steady or

[[Page 244]]

changing conditions of airspeed, engine power, and attitude.
    (c) If compliance with Sec. 23.145(b)(3) necessitates wing flap 
retraction to positions that are not fully retracted, the wing flap 
control lever settings corresponding to those positions must be 
positively located such that a definite change of direction of movement 
of the lever is necessary to select settings beyond those settings.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-49, 
61 FR 5165, Feb. 9, 1996]



Sec. 23.699  Wing flap position indicator.

    There must be a wing flap position indicator for--
    (a) Flap installations with only the retracted and fully extended 
position, unless--
    (1) A direct operating mechanism provides a sense of ``feel'' and 
position (such as when a mechanical linkage is employed); or
    (2) The flap position is readily determined without seriously 
detracting from other piloting duties under any flight condition, day or 
night; and
    (b) Flap installation with intermediate flap positions if--
    (1) Any flap position other than retracted or fully extended is used 
to show compliance with the performance requirements of this part; and
    (2) The flap installation does not meet the requirements of 
paragraph (a)(1) of this section.



Sec. 23.701  Flap interconnection.

    (a) The main wing flaps and related movable surfaces as a system 
must--
    (1) Be synchronized by a mechanical interconnection between the 
movable flap surfaces that is independent of the flap drive system; or 
by an approved equivalent means; or
    (2) Be designed so that the occurrence of any failure of the flap 
system that would result in an unsafe flight characteristic of the 
airplane is extremely improbable; or
    (b) The airplane must be shown to have safe flight characteristics 
with any combination of extreme positions of individual movable surfaces 
(mechanically interconnected surfaces are to be considered as a single 
surface).
    (c) If an interconnection is used in multiengine airplanes, it must 
be designed to account for the unsummetrical loads resulting from flight 
with the engines on one side of the plane of symmetry inoperative and 
the remaining engines at takeoff power. For single-engine airplanes, and 
multiengine airplanes with no slipstream effects on the flaps, it may be 
assumed that 100 percent of the critical air load acts on one side and 
70 percent on the other.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31821, Nov. 19, 1973; Amdt. 23-42, 56 FR 353, Jan. 3, 1991; 56 FR 
5455, Feb. 11, 1991; Amdt. 23-49, 61 FR 5165, Feb. 9, 1996]



Sec. 23.703  Takeoff warning system.

    For all airplanes with a maximum weight more than 6,000 pounds and 
all jets, unless it can be shown that a lift or longitudinal trim device 
that affects the takeoff performance of the airplane would not give an 
unsafe takeoff configuration when selected out of an approved takeoff 
position, a takeoff warning system must be installed and meet the 
following requirements:
    (a) The system must provide to the pilots an aural warning that is 
automatically activated during the initial portion of the takeoff role 
if the airplane is in a configuration that would not allow a safe 
takeoff. The warning must continue until--
    (1) The configuration is changed to allow safe takeoff, or
    (2) Action is taken by the pilot to abandon the takeoff roll.
    (b) The means used to activate the system must function properly for 
all authorized takeoff power settings and procedures and throughout the 
ranges of takeoff weights, altitudes, and temperatures for which 
certification is requested.
    (c) For the purpose of this section, an unsafe takeoff configuration 
is the inability to rotate or the inability to prevent an immediate 
stall after rotation.

[Doc. No. 27806, 61 FR 5166, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75757, Dec. 2, 2011]

[[Page 245]]

                              Landing Gear



Sec. 23.721  General.

    For commuter category airplanes that have a passenger seating 
configuration, excluding pilot seats, of 10 or more, the following 
general requirements for the landing gear apply:
    (a) The main landing-gear system must be designed so that if it 
fails due to overloads during takeoff and landing (assuming the 
overloads to act in the upward and aft directions), the failure mode is 
not likely to cause the spillage of enough fuel from any part of the 
fuel system to consitute a fire hazard.
    (b) Each airplane must be designed so that, with the airplane under 
control, it can be landed on a paved runway with any one or more 
landing-gear legs not extended without sustaining a structural component 
failure that is likely to cause the spillage of enough fuel to consitute 
a fire hazard.
    (c) Compliance with the provisions of this section may be shown by 
analysis or tests, or both.

[Amdt. 23-34, 52 FR 1830, Jan. 15, 1987]



Sec. 23.723  Shock absorption tests.

    (a) It must be shown that the limit load factors selected for design 
in accordance with Sec. 23.473 for takeoff and landing weights, 
respectively, will not be exceeded. This must be shown by energy 
absorption tests except that analysis based on tests conducted on a 
landing gear system with identical energy absorption characteristics may 
be used for increases in previously approved takeoff and landing 
weights.
    (b) The landing gear may not fail, but may yield, in a test showing 
its reserve energy absorption capacity, simulating a descent velocity of 
1.2 times the limit descent velocity, assuming wing lift equal to the 
weight of the airplane.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-23, 43 FR 50593, Oct. 30, 1978; Amdt. 23-49, 61 FR 
5166, Feb. 9, 1996]



Sec. 23.725  Limit drop tests.

    (a) If compliance with Sec. 23.723(a) is shown by free drop tests, 
these tests must be made on the complete airplane, or on units 
consisting of wheel, tire, and shock absorber, in their proper relation, 
from free drop heights not less than those determined by the following 
formula:

h (inches)=3.6 (W/S) \1/2\


However, the free drop height may not be less than 9.2 inches and need 
not be more than 18.7 inches.

    (b) If the effect of wing lift is provided for in free drop tests, 
the landing gear must be dropped with an effective weight equal to
[GRAPHIC] [TIFF OMITTED] TC28SE91.014

where--

We=the effective weight to be used in the drop test (lbs.);
h=specified free drop height (inches);
d=deflection under impact of the tire (at the approved inflation 
          pressure) plus the vertical component of the axle travel 
          relative to the drop mass (inches);
W=WM for main gear units (lbs), equal to the static weight on that unit 
          with the airplane in the level attitude (with the nose wheel 
          clear in the case of nose wheel type airplanes);
W=WT for tail gear units (lbs.), equal to the static weight on the tail 
          unit with the airplane in the tail-down attitude;
W=WN for nose wheel units lbs.), equal to the vertical component of the 
          static reaction that would exist at the nose wheel, assuming 
          that the mass of the airplane acts at the center of gravity 
          and exerts a force of 1.0 g downward and 0.33 g forward; and
L= the ratio of the assumed wing lift to the airplane weight, but not 
          more than 0.667.

    (c) The limit inertia load factor must be determined in a rational 
or conservative manner, during the drop test, using a landing gear unit 
attitude, and applied drag loads, that represent the landing conditions.
    (d) The value of d used in the computation of We in paragraph (b) of 
this section may not exceed the value actually obtained in the drop 
test.
    (e) The limit inertia load factor must be determined from the drop 
test in paragraph (b) of this section according to the following 
formula:
[GRAPHIC] [TIFF OMITTED] TC28SE91.015

where--


[[Page 246]]


nj=the load factor developed in the drop test (that is, the acceleration 
          (dv/dt) in gs recorded in the drop test) plus 1.0; and
We, W, and L are the same as in the drop test computation.

    (f) The value of n determined in accordance with paragraph (e) may 
not be more than the limit inertia load factor used in the landing 
conditions in Sec. 23.473.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13091, Aug. 13, 1969; Amdt. 23-48, 61 FR 5148, Feb. 9, 1996]



Sec. 23.726  Ground load dynamic tests.

    (a) If compliance with the ground load requirements of Sec. Sec. 
23.479 through 23.483 is shown dynamically by drop test, one drop test 
must be conducted that meets Sec. 23.725 except that the drop height 
must be--
    (1) 2.25 times the drop height prescribed in Sec. 23.725(a); or
    (2) Sufficient to develop 1.5 times the limit load factor.
    (b) The critical landing condition for each of the design conditions 
specified in Sec. Sec. 23.479 through 23.483 must be used for proof of 
strength.

[Amdt. 23-7, 34 FR 13091, Aug. 13, 1969]



Sec. 23.727  Reserve energy absorption drop test.

    (a) If compliance with the reserve energy absorption requirement in 
Sec. 23.723(b) is shown by free drop tests, the drop height may not be 
less than 1.44 times that specified in Sec. 23.725.
    (b) If the effect of wing lift is provided for, the units must be 
dropped with an effective mass equal to We=Wh/(h+d), when the symbols 
and other details are the same as in Sec. 23.725.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13091, Aug. 13, 1969]



Sec. 23.729  Landing gear extension and retraction system.

    (a) General. For airplanes with retractable landing gear, the 
following apply:
    (1) Each landing gear retracting mechanism and its supporting 
structure must be designed for maximum flight load factors with the gear 
retracted and must be designed for the combination of friction, inertia, 
brake torque, and air loads, occurring during retraction at any airspeed 
up to 1.6 VS1 with flaps retracted, and for any load factor 
up to those specified in Sec. 23.345 for the flaps-extended condition.
    (2) The landing gear and retracting mechanism, including the wheel 
well doors, must withstand flight loads, including loads resulting from 
all yawing conditions specified in Sec. 23.351, with the landing gear 
extended at any speed up to at least 1.6 VS1 with the flaps 
retracted.
    (b) Landing gear lock. There must be positive means (other than the 
use of hydraulic pressure) to keep the landing gear extended.
    (c) Emergency operation. For a landplane having retractable landing 
gear that cannot be extended manually, there must be means to extend the 
landing gear in the event of either--
    (1) Any reasonably probable failure in the normal landing gear 
operation system; or
    (2) Any reasonably probable failure in a power source that would 
prevent the operation of the normal landing gear operation system.
    (d) Operation test. The proper functioning of the retracting 
mechanism must be shown by operation tests.
    (e) Position indicator. If a retractable landing gear is used, there 
must be a landing gear position indicator (as well as necessary switches 
to actuate the indicator) or other means to inform the pilot that each 
gear is secured in the extended (or retracted) position. If switches are 
used, they must be located and coupled to the landing gear mechanical 
system in a manner that prevents an erroneous indication of either 
``down and locked'' if each gear is not in the fully extended position, 
or ``up and locked'' if each landing gear is not in the fully retracted 
position.
    (f) Landing gear warning. For landplanes, the following aural or 
equally effective landing gear warning devices must be provided:
    (1) A device that functions continuously when one or more throttles 
are closed beyond the power settings normally used for landing approach 
if the landing gear is not fully extended and locked. A throttle stop 
may not be used in place of an aural device. If

[[Page 247]]

there is a manual shutoff for the warning device prescribed in this 
paragraph, the warning system must be designed so that when the warning 
has been suspended after one or more throttles are closed, subsequent 
retardation of any throttle to, or beyond, the position for normal 
landing approach will activate the warning device.
    (2) A device that functions continuously when the wing flaps are 
extended beyond the maximum approach flap position, using a normal 
landing procedure, if the landing gear is not fully extended and locked. 
There may not be a manual shutoff for this warning device. The flap 
position sensing unit may be installed at any suitable location. The 
system for this device may use any part of the system (including the 
aural warning device) for the device required in paragraph (f)(1) of 
this section.
    (g) Equipment located in the landing gear bay. If the landing gear 
bay is used as the location for equipment other than the landing gear, 
that equipment must be designed and installed to minimize damage from 
items such as a tire burst, or rocks, water, and slush that may enter 
the landing gear bay.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13091, Aug. 13, 1969; Amdt. 23-21, 43 FR 2318, Jan. 1978; Amdt. 23-
26, 45 FR 60171, Sept. 11, 1980; Amdt. 23-45, 58 FR 42164, Aug. 6, 1993; 
Amdt. 23-49, 61 FR 5166, Feb. 9, 1996]



Sec. 23.731  Wheels.

    (a) The maximum static load rating of each wheel may not be less 
than the corresponding static ground reaction with--
    (1) Design maximum weight; and
    (2) Critical center of gravity.
    (b) The maximum limit load rating of each wheel must equal or exceed 
the maximum radial limit load determined under the applicable ground 
load requirements of this part.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42165, Aug. 6, 1993]



Sec. 23.733  Tires.

    (a) Each landing gear wheel must have a tire whose approved tire 
ratings (static and dynamic) are not exceeded--
    (1) By a load on each main wheel tire) to be compared to the static 
rating approved for such tires) equal to the corresponding static ground 
reaction under the design maximum weight and critical center of gravity; 
and
    (2) By a load on nose wheel tires (to be compared with the dynamic 
rating approved for such tires) equal to the reaction obtained at the 
nose wheel, assuming the mass of the airplane to be concentrated at the 
most critical center of gravity and exerting a force of 1.0 W downward 
and 0.31 W forward (where W is the design maximum weight), with the 
reactions distributed to the nose and main wheels by the principles of 
statics and with the drag reaction at the ground applied only at wheels 
with brakes.
    (b) If specially constructed tires are used, the wheels must be 
plainly and conspicuously marked to that effect. The markings must 
include the make, size, number of plies, and identification marking of 
the proper tire.
    (c) Each tire installed on a retractable landing gear system must, 
at the maximum size of the tire type expected in service, have a 
clearance to surrounding structure and systems that is adequate to 
prevent contact between the tire and any part of the structure of 
systems.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13092, Aug. 13, 1969; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. 
23-45, 58 FR 42165, Aug. 6, 1993]



Sec. 23.735  Brakes.

    (a) Brakes must be provided. The landing brake kinetic energy 
capacity rating of each main wheel brake assembly must not be less than 
the kinetic energy absorption requirements determined under either of 
the following methods:
    (1) The brake kinetic energy absorption requirements must be based 
on a conservative rational analysis of the sequence of events expected 
during landing at the design landing weight.
    (2) Instead of a rational analysis, the kinetic energy absorption 
requirements for each main wheel brake assembly may be derived from the 
following formula:

KE=0.0443 WV\2\/N


[[Page 248]]


where--

KE=Kinetic energy per wheel (ft.-lb.);
W=Design landing weight (lb.);
V=Airplane speed in knots. V must be not less than VS[radic], 
          the poweroff stalling speed of the airplane at sea level, at 
          the design landing weight, and in the landing configuration; 
          and
N=Number of main wheels with brakes.

    (b) Brakes must be able to prevent the wheels from rolling on a 
paved runway with takeoff power on the critical engine, but need not 
prevent movement of the airplane with wheels locked.
    (c) During the landing distance determination required by Sec. 
23.75, the pressure on the wheel braking system must not exceed the 
pressure specified by the brake manufacturer.
    (d) If antiskid devices are installed, the devices and associated 
systems must be designed so that no single probable malfunction or 
failure will result in a hazardous loss of braking ability or 
directional control of the airplane.
    (e) For airplanes required to meet Sec. 23.55, the rejected takeoff 
brake kinetic energy capacity rating of each main wheel brake assembly 
may not be less than the kinetic energy absorption requirements 
determined under either of the following methods--
    (1) The brake kinetic energy absorption requirements must be based 
on a conservative rational analysis of the sequence of events expected 
during a rejected takeoff at the design takeoff weight.
    (2) Instead of a rational analysis, the kinetic energy absorption 
requirements for each main wheel brake assembly may be derived from the 
following formula--

KE = 0.0443 WV\2\/N where;
KE = Kinetic energy per wheel (ft.-lbs.);
W = Design takeoff weight (lbs.);
V = Ground speed, in knots, associated with the maximum value of 
V1 selected in accordance with Sec. 23.51(c)(1);
N = Number of main wheels with brakes.

[Amdt. 23-7, 34 FR 13092, Aug. 13, 1969, as amended by Amdt. 23-24, 44 
FR 68742, Nov. 29, 1979; Amdt. 23-42, 56 FR 354, Jan. 3, 1991; Amdt. 23-
49, 61 FR 5166, Feb. 9, 1996; Amdt. 23-62, 76 FR 75757, Dec. 2, 2011]



Sec. 23.737  Skis.

    The maximum limit load rating for each ski must equal or exceed the 
maximum limit load determined under the applicable ground load 
requirements of this part.

[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]



Sec. 23.745  Nose/tail wheel steering.

    (a) If nose/tail wheel steering is installed, it must be 
demonstrated that its use does not require exceptional pilot skill 
during takeoff and landing, in crosswinds, or in the event of an engine 
failure; or its use must be limited to low speed maneuvering.
    (b) Movement of the pilot's steering control must not interfere with 
the retraction or extension of the landing gear.

[Doc. No. 27806, 61 FR 5166, Feb. 9, 1996]

                            Floats and Hulls



Sec. 23.751  Main float buoyancy.

    (a) Each main float must have--
    (1) A buoyancy of 80 percent in excess of the buoyancy required by 
that float to support its portion of the maximum weight of the seaplane 
or amphibian in fresh water; and
    (2) Enough watertight compartments to provide reasonable assurance 
that the seaplane or amphibian will stay afloat without capsizing if any 
two compartments of any main float are flooded.
    (b) Each main float must contain at least four watertight 
compartments approximately equal in volume.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42165, Aug. 6, 1993]



Sec. 23.753  Main float design.

    Each seaplane main float must meet the requirements of Sec. 23.521.

[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]



Sec. 23.755  Hulls.

    (a) The hull of a hull seaplane or amphibian of 1,500 pounds or more 
maximum weight must have watertight compartments designed and arranged 
so that the hull auxiliary floats, and tires (if used), will keep the 
airplane

[[Page 249]]

afloat without capsizing in fresh water when--
    (1) For airplanes of 5,000 pounds or more maximum weight, any two 
adjacent compartments are flooded; and
    (2) For airplanes of 1,500 pounds up to, but not including, 5,000 
pounds maximum weight, any single compartment is flooded.
    (b) Watertight doors in bulkheads may be used for communication 
between compartments.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42165, Aug. 6, 1993; Amdt. 23-48, 61 FR 5148, Feb. 9, 1996]



Sec. 23.757  Auxiliary floats.

    Auxiliary floats must be arranged so that, when completely submerged 
in fresh water, they provide a righting moment of at least 1.5 times the 
upsetting moment caused by the seaplane or amphibian being tilted.

                   Personnel and Cargo Accommodations



Sec. 23.771  Pilot compartment.

    For each pilot compartment--
    (a) The compartment and its equipment must allow each pilot to 
perform his duties without unreasonable concentration or fatigue;
    (b) Where the flight crew are separated from the passengers by a 
partition, an opening or openable window or door must be provided to 
facilitate communication between flight crew and the passengers; and
    (c) The aerodynamic controls listed in Sec. 23.779, excluding 
cables and control rods, must be located with respect to the propellers 
so that no part of the pilot or the controls lies in the region between 
the plane of rotation of any inboard propeller and the surface generated 
by a line passing through the center of the propeller hub making an 
angle of 5 degrees forward or aft of the plane of rotation of the 
propeller.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31821, Nov. 19, 1973]



Sec. 23.773  Pilot compartment view.

    (a) Each pilot compartment must be--
    (1) Arranged with sufficiently extensive, clear and undistorted view 
to enable the pilot to safely taxi, takeoff, approach, land, and perform 
any maneuvers within the operating limitations of the airplane.
    (2) Free from glare and reflections that could interfere with the 
pilot's vision. Compliance must be shown in all operations for which 
certification is requested; and
    (3) Designed so that each pilot is protected from the elements so 
that moderate rain conditions do not unduly impair the pilot's view of 
the flight path in normal flight and while landing.
    (b) Each pilot compartment must have a means to either remove or 
prevent the formation of fog or frost on an area of the internal portion 
of the windshield and side windows sufficiently large to provide the 
view specified in paragraph (a)(1) of this section. Compliance must be 
shown under all expected external and internal ambient operating 
conditions, unless it can be shown that the windshield and side windows 
can be easily cleared by the pilot without interruption of normal pilot 
duties.

[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993; 71 FR 537, Jan. 5, 2006]



Sec. 23.775  Windshields and windows.

    (a) The internal panels of windshields and windows must be 
constructed of a nonsplintering material, such as nonsplintering safety 
glass.
    (b) The design of windshields, windows, and canopies in pressurized 
airplanes must be based on factors peculiar to high altitude operation, 
including--
    (1) The effects of continuous and cyclic pressurization loadings;
    (2) The inherent characteristics of the material used; and
    (3) The effects of temperatures and temperature gradients.
    (c) On pressurized airplanes, if certification for operation up to 
and including 25,000 feet is requested, an enclosure canopy including a 
representative part of the installation must be subjected to special 
tests to account for the combined effects of continuous and cyclic 
pressurization loadings and flight loads, or compliance with the fail-
safe requirements of paragraph (d) of this section must be shown.

[[Page 250]]

    (d) If certification for operation above 25,000 feet is requested 
the windshields, window panels, and canopies must be strong enough to 
withstand the maximum cabin pressure differential loads combined with 
critical aerodynamic pressure and temperature effects, after failure of 
any load-carrying element of the windshield, window panel, or canopy.
    (e) The windshield and side windows forward of the pilot's back when 
the pilot is seated in the normal flight position must have a luminous 
transmittance value of not less than 70 percent.
    (f) Unless operation in known or forecast icing conditions is 
prohibited by operating limitations, a means must be provided to prevent 
or to clear accumulations of ice from the windshield so that the pilot 
has adequate view for taxi, takeoff, approach, landing, and to perform 
any maneuvers within the operating limitations of the airplane.
    (g) In the event of any probable single failure, a transparency 
heating system must be incapable of raising the temperature of any 
windshield or window to a point where there would be--
    (1) Structural failure that adversely affects the integrity of the 
cabin; or
    (2) There would be a danger of fire.
    (h) In addition, for commuter category airplanes, the following 
applies:
    (1) Windshield panes directly in front of the pilots in the normal 
conduct of their duties, and the supporting structures for these panes, 
must withstand, without penetration, the impact of a two-pound bird when 
the velocity of the airplane (relative to the bird along the airplane's 
flight path) is equal to the airplane's maximum approach flap speed.
    (2) The windshield panels in front of the pilots must be arranged so 
that, assuming the loss of vision through any one panel, one or more 
panels remain available for use by a pilot seated at a pilot station to 
permit continued safe flight and landing.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13092, Aug. 13, 1969; Amdt. 23-45, 58 FR 42165, Aug. 6, 1993; 58 FR 
51970, Oct. 5, 1993; Amdt. 23-49, 61 FR 5166, Feb. 9, 1996]



Sec. 23.777  Cockpit controls.

    (a) Each cockpit control must be located and (except where its 
function is obvious) identified to provide convenient operation and to 
prevent confusion and inadvertent operation.
    (b) The controls must be located and arranged so that the pilot, 
when seated, has full and unrestricted movement of each control without 
interference from either his clothing or the cockpit structure.
    (c) Powerplant controls must be located--
    (1) For multiengine airplanes, on the pedestal or overhead at or 
near the center of the cockpit;
    (2) For single and tandem seated single-engine airplanes, on the 
left side console or instrument panel;
    (3) For other single-engine airplanes at or near the center of the 
cockpit, on the pedestal, instrument panel, or overhead; and
    (4) For airplanes, with side-by-side pilot seats and with two sets 
of powerplant controls, on left and right consoles.
    (d) When separate and distinct control levers are co-located (such 
as located together on the pedestal), the control location order from 
left to right must be power (thrust) lever, propeller (rpm control), and 
mixture control (condition lever and fuel cut-off for turbine-powered 
airplanes). Power (thrust) levers must be easily distinguishable from 
other controls, and provide for accurate, consistent operation. 
Carburetor heat or alternate air control must be to the left of the 
throttle or at least eight inches from the mixture control when located 
other than on a pedestal. Carburetor heat or alternate air control, when 
located on a pedestal, must be aft or below the power (thrust) lever. 
Supercharger controls must be located below or aft of the propeller 
controls. Airplanes with tandem seating or single-place airplanes may 
utilize control locations on the left side of the cabin compartment; 
however, location order from left to right must be power (thrust) lever, 
propeller (rpm control), and mixture control.
    (e) Identical powerplant controls for each engine must be located to 
prevent

[[Page 251]]

confusion as to the engines they control.
    (1) Conventional multiengine powerplant controls must be located so 
that the left control(s) operates the left engines(s) and the right 
control(s) operates the right engine(s).
    (2) On twin-engine airplanes with front and rear engine locations 
(tandem), the left powerplant controls must operate the front engine and 
the right powerplant controls must operate the rear engine.
    (f) Wing flap and auxiliary lift device controls must be located--
    (1) Centrally, or to the right of the pedestal or powerplant 
throttle control centerline; and
    (2) Far enough away from the landing gear control to avoid 
confusion.
    (g) The landing gear control must be located to the left of the 
throttle centerline or pedestal centerline.
    (h) Each fuel feed selector control must comply with Sec. 23.995 
and be located and arranged so that the pilot can see and reach it 
without moving any seat or primary flight control when his seat is at 
any position in which it can be placed.
    (1) For a mechanical fuel selector:
    (i) The indication of the selected fuel valve position must be by 
means of a pointer and must provide positive identification and feel 
(detent, etc.) of the selected position.
    (ii) The position indicator pointer must be located at the part of 
the handle that is the maximum dimension of the handle measured from the 
center of rotation.
    (2) For electrical or electronic fuel selector:
    (i) Digital controls or electrical switches must be properly 
labelled.
    (ii) Means must be provided to indicate to the flight crew the tank 
or function selected. Selector switch position is not acceptable as a 
means of indication. The ``off'' or ``closed'' position must be 
indicated in red.
    (3) If the fuel valve selector handle or electrical or digital 
selection is also a fuel shut-off selector, the off position marking 
must be colored red. If a separate emergency shut-off means is provided, 
it also must be colored red.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13092, Aug. 13, 1969; Amdt. 23-33, 51 FR 26656, July 24, 1986; Amdt. 
23-51, 61 FR 5136, Feb. 9, 1996; Amdt. 23-62, 76 FR 75757, Dec. 2, 2011]



Sec. 23.779  Motion and effect of cockpit controls.

    Cockpit controls must be designed so that they operate in accordance 
with the following movement and actuation:
    (a) Aerodynamic controls:

 
                                                Motion and effect
 
(1) Primary controls:
    Aileron...........................  Right (clockwise) for right wing
                                         down.
    Elevator..........................  Rearward for nose up.
    Rudder............................  Right pedal forward for nose
                                         right.
(2) Secondary controls:
    Flaps (or auxiliary lift devices).  Forward or up for flaps up or
                                         auxiliary device stowed;
                                         rearward or down for flaps down
                                         or auxiliary device deployed.
    Trim tabs (or equivalent).........  Switch motion or mechanical
                                         rotation of control to produce
                                         similar rotation of the
                                         airplane about an axis parallel
                                         to the axis control. Axis of
                                         roll trim control may be
                                         displaced to accommodate
                                         comfortable actuation by the
                                         pilot. For single-engine
                                         airplanes, direction of pilot's
                                         hand movement must be in the
                                         same sense as airplane response
                                         for rudder trim if only a
                                         portion of a rotational element
                                         is accessible.
 

    (b) Powerplant and auxiliary controls:

 
                                                Motion and effect
 
(1) Powerplant controls:
    Power (thrust) lever..............  Forward to increase forward
                                         thrust and rearward to increase
                                         rearward thrust.
    Propellers........................  Forward to increase rpm.
    Mixture...........................  Forward or upward for rich.

[[Page 252]]

 
    Fuel..............................  Forward for open.
    Carburetor, air heat or alternate   Forward or upward for cold.
     air.
    Supercharger......................  Forward or upward for low
                                         blower.
    Turbosuper-chargers...............  Forward, upward, or clockwise to
                                         increase pressure.
    Rotary controls...................  Clockwise from off to full on.
(2) Auxiliary controls:
    Fuel tank selector................  Right for right tanks, left for
                                         left tanks.
    Landing gear......................  Down to extend.
    Speed brakes......................  Aft to extend.
 


[Amdt. 23-33, 51 FR 26656, July 24, 1986, as amended by Amdt. 23-51, 61 
FR 5136, Feb. 9, 1996]



Sec. 23.781  Cockpit control knob shape.

    (a) Flap and landing gear control knobs must conform to the general 
shapes (but not necessarily the exact sizes or specific proportions) in 
the following figure:
[GRAPHIC] [TIFF OMITTED] TC28SE91.016


[[Page 253]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.017

    (b) Powerplant control knobs must conform to the general shapes (but 
not necessarily the exact sizes or specific proportions) in the 
following figure:

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-33, 51 FR 26657, July 24, 1986]



Sec. 23.783  Doors.

    (a) Each closed cabin with passenger accommodations must have at 
least one adequate and easily accessible external door.
    (b) Passenger doors must not be located with respect to any 
propeller disk or any other potential hazard so as to endanger persons 
using the door.
    (c) Each external passenger or crew door must comply with the 
following requirements:
    (1) There must be a means to lock and safeguard the door against 
inadvertent opening during flight by persons, by cargo, or as a result 
of mechanical failure.
    (2) The door must be openable from the inside and the outside when 
the internal locking mechanism is in the locked position.

[[Page 254]]

    (3) There must be a means of opening which is simple and obvious and 
is arranged and marked inside and outside so that the door can be 
readily located, unlocked, and opened, even in darkness.
    (4) The door must meet the marking requirements of Sec. 23.811 of 
this part.
    (5) The door must be reasonably free from jamming as a result of 
fuselage deformation in an emergency landing.
    (6) Auxiliary locking devices that are actuated externally to the 
airplane may be used but such devices must be overridden by the normal 
internal opening means.
    (d) In addition, each external passenger or crew door, for a 
commuter category airplane, must comply with the following requirements:
    (1) Each door must be openable from both the inside and outside, 
even though persons may be crowded against the door on the inside of the 
airplane.
    (2) If inward opening doors are used, there must be a means to 
prevent occupants from crowding against the door to the extent that 
would interfere with opening the door.
    (3) Auxiliary locking devices may be used.
    (e) Each external door on a commuter category airplane, each 
external door forward of any engine or propeller on a normal, utility, 
or acrobatic category airplane, and each door of the pressure vessel on 
a pressurized airplane must comply with the following requirements:
    (1) There must be a means to lock and safeguard each external door, 
including cargo and service type doors, against inadvertent opening in 
flight, by persons, by cargo, or as a result of mechanical failure or 
failure of a single structural element, either during or after closure.
    (2) There must be a provision for direct visual inspection of the 
locking mechanism to determine if the external door, for which the 
initial opening movement is not inward, is fully closed and locked. The 
provisions must be discernible, under operating lighting conditions, by 
a crewmember using a flashlight or an equivalent lighting source.
    (3) There must be a visual warning means to signal a flight 
crewmember if the external door is not fully closed and locked. The 
means must be designed so that any failure, or combination of failures, 
that would result in an erroneous closed and locked indication is 
improbable for doors for which the initial opening movement is not 
inward.
    (f) In addition, for commuter category airplanes, the following 
requirements apply:
    (1) Each passenger entry door must qualify as a floor level 
emergency exit. This exit must have a rectangular opening of not less 
than 24 inches wide by 48 inches high, with corner radii not greater 
than one-third the width of the exit.
    (2) If an integral stair is installed at a passenger entry door, the 
stair must be designed so that, when subjected to the inertia loads 
resulting from the ultimate static load factors in Sec. 23.561(b)(2) 
and following the collapse of one or more legs of the landing gear, it 
will not reduce the effectiveness of emergency egress through the 
passenger entry door.
    (g) If lavatory doors are installed, they must be designed to 
preclude an occupant from becoming trapped inside the lavatory. If a 
locking mechanism is installed, it must be capable of being unlocked 
from outside of the lavatory.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-36, 53 FR 30813, Aug. 15, 1988; Amdt. 23-46, 59 FR 
25772, May 17, 1994; Amdt. 23-49, 61 FR 5166, Feb. 9, 1996]



Sec. 23.785  Seats, berths, litters, safety belts, and shoulder harnesses.

    There must be a seat or berth for each occupant that meets the 
following:
    (a) Each seat/restraint system and the supporting structure must be 
designed to support occupants weighing at least 215 pounds when 
subjected to the maximum load factors corresponding to the specified 
flight and ground load conditions, as defined in the approved operating 
envelope of the airplane. In addition, these loads must be multiplied by 
a factor of 1.33 in determining the strength of all fittings and the 
attachment of--

[[Page 255]]

    (1) Each seat to the structure; and
    (2) Each safety belt and shoulder harness to the seat or structure.
    (b) Each forward-facing or aft-facing seat/restraint system in 
normal, utility, or acrobatic category airplanes must consist of a seat, 
a safety belt, and a shoulder harness, with a metal-to-metal latching 
device, that are designed to provide the occupant protection provisions 
required in Sec. 23.562. Other seat orientations must provide the same 
level of occupant protection as a forward-facing or aft-facing seat with 
a safety belt and a shoulder harness, and must provide the protection 
provisions of Sec. 23.562.
    (c) For commuter category airplanes, each seat and the supporting 
structure must be designed for occupants weighing at least 170 pounds 
when subjected to the inertia loads resulting from the ultimate static 
load factors prescribed in Sec. 23.561(b)(2) of this part. Each 
occupant must be protected from serious head injury when subjected to 
the inertia loads resulting from these load factors by a safety belt and 
shoulder harness, with a metal-to-metal latching device, for the front 
seats and a safety belt, or a safety belt and shoulder harness, with a 
metal-to-metal latching device, for each seat other than the front 
seats.
    (d) Each restraint system must have a single-point release for 
occupant evacuation.
    (e) The restraint system for each crewmember must allow the 
crewmember, when seated with the safety belt and shoulder harness 
fastened, to perform all functions necessary for flight operations.
    (f) Each pilot seat must be designed for the reactions resulting 
from the application of pilot forces to the primary flight controls as 
prescribed in Sec. 23.395 of this part.
    (g) There must be a means to secure each safety belt and shoulder 
harness, when not in use, to prevent interference with the operation of 
the airplane and with rapid occupant egress in an emergency.
    (h) Unless otherwise placarded, each seat in a utility or acrobatic 
category airplane must be designed to accommodate an occupant wearing a 
parachute.
    (i) The cabin area surrounding each seat, including the structure, 
interior walls, instrument panel, control wheel, pedals, and seats 
within striking distance of the occupant's head or torso (with the 
restraint system fastened) must be free of potentially injurious 
objects, sharp edges, protuberances, and hard surfaces. If energy 
absorbing designs or devices are used to meet this requirement, they 
must protect the occupant from serious injury when the occupant is 
subjected to the inertia loads resulting from the ultimate static load 
factors prescribed in Sec. 23.561(b)(2) of this part, or they must 
comply with the occupant protection provisions of Sec. 23.562 of this 
part, as required in paragraphs (b) and (c) of this section.
    (j) Each seat track must be fitted with stops to prevent the seat 
from sliding off the track.
    (k) Each seat/restraint system may use design features, such as 
crushing or separation of certain components, to reduce occupant loads 
when showing compliance with the requirements of Sec. 23.562 of this 
part; otherwise, the system must remain intact.
    (l) For the purposes of this section, a front seat is a seat located 
at a flight crewmember station or any seat located alongside such a 
seat.
    (m) Each berth, or provisions for a litter, installed parallel to 
the longitudinal axis of the airplane, must be designed so that the 
forward part has a padded end-board, canvas diaphragm, or equivalent 
means that can withstand the load reactions from a 215-pound occupant 
when subjected to the inertia loads resulting from the ultimate static 
load factors of Sec. 23.561(b)(2) of this part. In addition--
    (1) Each berth or litter must have an occupant restraint system and 
may not have corners or other parts likely to cause serious injury to a 
person occupying it during emergency landing conditions; and
    (2) Occupant restraint system attachments for the berth or litter 
must withstand the inertia loads resulting from the ultimate static load 
factors of Sec. 23.561(b)(2) of this part.
    (n) Proof of compliance with the static strength requirements of 
this section for seats and berths approved as

[[Page 256]]

part of the type design and for seat and berth installations may be 
shown by--
    (1) Structural analysis, if the structure conforms to conventional 
airplane types for which existing methods of analysis are known to be 
reliable;
    (2) A combination of structural analysis and static load tests to 
limit load; or
    (3) Static load tests to ultimate loads.

[Amdt. 23-36, 53 FR 30813, Aug. 15, 1988; Amdt. 23-36, 54 FR 50737, Dec. 
11, 1989; Amdt. 23-49, 61 FR 5167, Feb. 9, 1996]



Sec. 23.787  Baggage and cargo compartments.

    (a) Each baggage and cargo compartment must:
    (1) Be designed for its placarded maximum weight of contents and for 
the critical load distributions at the appropriate maximum load factors 
corresponding to the flight and ground load conditions of this part.
    (2) Have means to prevent the contents of any compartment from 
becoming a hazard by shifting, and to protect any controls, wiring, 
lines, equipment or accessories whose damage or failure would affect 
safe operations.
    (3) Have a means to protect occupants from injury by the contents of 
any compartment, located aft of the occupants and separated by 
structure, when the ultimate forward inertial load factor is 9g and 
assuming the maximum allowed baggage or cargo weight for the 
compartment.
    (b) Designs that provide for baggage or cargo to be carried in the 
same compartment as passengers must have a means to protect the 
occupants from injury when the baggage or cargo is subjected to the 
inertial loads resulting from the ultimate static load factors of Sec. 
23.561(b)(3), assuming the maximum allowed baggage or cargo weight for 
the compartment.
    (c) For airplanes that are used only for the carriage of cargo, the 
flightcrew emergency exits must meet the requirements of Sec. 23.807 
under any cargo loading conditions.

[Doc. No. 27806, 61 FR 5167, Feb. 9, 1996]



Sec. 23.791  Passenger information signs.

    For those airplanes in which the flightcrew members cannot observe 
the other occupants' seats or where the flightcrew members' compartment 
is separated from the passenger compartment, there must be at least one 
illuminated sign (using either letters or symbols) notifying all 
passengers when seat belts should be fastened. Signs that notify when 
seat belts should be fastened must:
    (a) When illuminated, be legible to each person seated in the 
passenger compartment under all probable lighting conditions; and
    (b) Be installed so that a flightcrew member can, when seated at the 
flightcrew member's station, turn the illumination on and off.

[Doc. No. 27806, 61 FR 5167, Feb. 9, 1996]



Sec. 23.803  Emergency evacuation.

    (a) For commuter category airplanes, an evacuation demonstration 
must be conducted utilizing the maximum number of occupants for which 
certification is desired. The demonstration must be conducted under 
simulated night conditions using only the emergency exits on the most 
critical side of the airplane. The participants must be representative 
of average airline passengers with no prior practice or rehearsal for 
the demonstration. Evacuation must be completed within 90 seconds.
    (b) In addition, when certification to the emergency exit provisions 
of Sec. 23.807(d)(4) is requested, only the emergency lighting system 
required by Sec. 23.812 may be used to provide cabin interior 
illumination during the evacuation demonstration required in paragraph 
(a) of this section.

[Amdt. 23-34, 52 FR 1831, Jan. 15, 1987, as amended by Amdt. 23-46, 59 
FR 25773, May 17, 1994]



Sec. 23.805  Flightcrew emergency exits.

    For airplanes where the proximity of the passenger emergency exits 
to the flightcrew area does not offer a convenient and readily 
accessible means of evacuation for the flightcrew, the following apply:
    (a) There must be either one emergency exit on each side of the 
airplane, or a top hatch emergency exit, in the flightcrew area;

[[Page 257]]

    (b) Each emergency exit must be located to allow rapid evacuation of 
the crew and have a size and shape of at least a 19- by 20-inch 
unobstructed rectangular opening; and
    (c) For each emergency exit that is not less than six feet from the 
ground, an assisting means must be provided. The assisting means may be 
a rope or any other means demonstrated to be suitable for the purpose. 
If the assisting means is a rope, or an approved device equivalent to a 
rope, it must be--
    (1) Attached to the fuselage structure at or above the top of the 
emergency exit opening or, for a device at a pilot's emergency exit 
window, at another approved location if the stowed device, or its 
attachment, would reduce the pilot's view; and
    (2) Able (with its attachment) to withstand a 400-pound static load.

[Doc. No. 26324, 59 FR 25773, May 17, 1994]



Sec. 23.807  Emergency exits.

    (a) Number and location. Emergency exits must be located to allow 
escape without crowding in any probable crash attitude. The airplane 
must have at least the following emergency exits:
    (1) For all airplanes with a seating capacity of two or more, 
excluding airplanes with canopies, at least one emergency exit on the 
opposite side of the cabin from the main door specified in Sec. 23.783 
of this part.
    (2) [Reserved]
    (3) If the pilot compartment is separated from the cabin by a door 
that is likely to block the pilot's escape in a minor crash, there must 
be an exit in the pilot's compartment. The number of exits required by 
paragraph (a)(1) of this section must then be separately determined for 
the passenger compartment, using the seating capacity of that 
compartment.
    (4) Emergency exits must not be located with respect to any 
propeller disk or any other potential hazard so as to endanger persons 
using that exit.
    (b) Type and operation. Emergency exits must be movable windows, 
panels, canopies, or external doors, openable from both inside and 
outside the airplane, that provide a clear and unobstructed opening 
large enough to admit a 19-by-26-inch ellipse. Auxiliary locking devices 
used to secure the airplane must be designed to be overridden by the 
normal internal opening means. The inside handles of emergency exits 
that open outward must be adequately protected against inadvertent 
operation. In addition, each emergency exit must--
    (1) Be readily accessible, requiring no exceptional agility to be 
used in emergencies;
    (2) Have a method of opening that is simple and obvious;
    (3) Be arranged and marked for easy location and operation, even in 
darkness;
    (4) Have reasonable provisions against jamming by fuselage 
deformation; and
    (5) In the case of acrobatic category airplanes, allow each occupant 
to abandon the airplane at any speed between VSO and 
VD; and
    (6) In the case of utility category airplanes certificated for 
spinning, allow each occupant to abandon the airplane at the highest 
speed likely to be achieved in the maneuver for which the airplane is 
certificated.
    (c) Tests. The proper functioning of each emergency exit must be 
shown by tests.
    (d) Doors and exits. In addition, for commuter category airplanes, 
the following requirements apply:
    (1) In addition to the passenger entry door--
    (i) For an airplane with a total passenger seating capacity of 15 or 
fewer, an emergency exit, as defined in paragraph (b) of this section, 
is required on each side of the cabin; and
    (ii) For an airplane with a total passenger seating capacity of 16 
through 19, three emergency exits, as defined in paragraph (b) of this 
section, are required with one on the same side as the passenger entry 
door and two on the side opposite the door.
    (2) A means must be provided to lock each emergency exit and to 
safeguard against its opening in flight, either inadvertently by persons 
or as a result of mechanical failure. In addition, a means for direct 
visual inspection of the locking mechanism must be provided to determine 
that each emergency exit for which the initial opening movement is 
outward is fully locked.

[[Page 258]]

    (3) Each required emergency exit, except floor level exits, must be 
located over the wing or, if not less than six feet from the ground, 
must be provided with an acceptable means to assist the occupants to 
descend to the ground. Emergency exits must be distributed as uniformly 
as practical, taking into account passenger seating configuration.
    (4) Unless the applicant has complied with paragraph (d)(1) of this 
section, there must be an emergency exit on the side of the cabin 
opposite the passenger entry door, provided that--
    (i) For an airplane having a passenger seating configuration of nine 
or fewer, the emergency exit has a rectangular opening measuring not 
less than 19 inches by 26 inches high with corner radii not greater than 
one-third the width of the exit, located over the wing, with a step up 
inside the airplane of not more than 29 inches and a step down outside 
the airplane of not more than 36 inches;
    (ii) For an airplane having a passenger seating configuration of 10 
to 19 passengers, the emergency exit has a rectangular opening measuring 
not less than 20 inches wide by 36 inches high, with corner radii not 
greater than one-third the width of the exit, and with a step up inside 
the airplane of not more than 20 inches. If the exit is located over the 
wing, the step down outside the airplane may not exceed 27 inches; and
    (iii) The airplane complies with the additional requirements of 
Sec. Sec. 23.561(b)(2)(iv), 23.803(b), 23.811(c), 23.812, 23.813(b), 
and 23.815.
    (e) For multiengine airplanes, ditching emergency exits must be 
provided in accordance with the following requirements, unless the 
emergency exits required by paragraph (a) or (d) of this section already 
comply with them:
    (1) One exit above the waterline on each side of the airplane having 
the dimensions specified in paragraph (b) or (d) of this section, as 
applicable; and
    (2) If side exits cannot be above the waterline, there must be a 
readily accessible overhead hatch emergency exit that has a rectangular 
opening measuring not less than 20 inches wide by 36 inches long, with 
corner radii not greater than one-third the width of the exit.
    (3) In lieu of paragraph (e)(2) of this section, if any side exit(s) 
cannot be above the waterline, a device may be placed at each of such 
exit(s) prior to ditching. This device must slow the inflow of water 
when such exit(s) is opened with the airplane ditched. For commuter 
category airplanes, the clear opening of such exit(s) must meet the 
requirements defined in paragraph (d) of this section.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13092, Aug. 13, 1969; Amdt. 23-10, 36 FR 2864, Feb. 11, 1971; Amdt. 
23-34, 52 FR 1831, Jan. 15, 1987; Amdt. 23-36, 53 FR 30814, Aug. 15, 
1988; 53 FR 34194, Sept. 2, 1988; Amdt. 23-46, 59 FR 25773, May 17, 
1994; Amdt. 23-49, 61 FR 5167, Feb. 9, 1996; Amdt. 23-62, 76 FR 75757, 
Dec. 2, 2011]



Sec. 23.811  Emergency exit marking.

    (a) Each emergency exit and external door in the passenger 
compartment must be externally marked and readily identifiable from 
outside the airplane by--
    (1) A conspicuous visual identification scheme; and
    (2) A permanent decal or placard on or adjacent to the emergency 
exit which shows the means of opening the emergency exit, including any 
special instructions, if applicable.
    (b) In addition, for commuter category airplanes, these exits and 
doors must be internally marked with the word ``exit'' by a sign which 
has white letters 1 inch high on a red background 2 inches high, be 
self-illuminated or independently, internally electrically illuminated, 
and have a minimum brightness of at least 160 microlamberts. The color 
may be reversed if the passenger compartment illumination is essentially 
the same.
    (c) In addition, when certification to the emergency exit provisions 
of Sec. 23.807(d)(4) is requested, the following apply:
    (1) Each emergency exit, its means of access, and its means of 
opening, must be conspicuously marked;
    (2) The identity and location of each emergency exit must be 
recognizable from a distance equal to the width of the cabin;
    (3) Means must be provided to assist occupants in locating the 
emergency exits in conditions of dense smoke;

[[Page 259]]

    (4) The location of the operating handle and instructions for 
opening each emergency exit from inside the airplane must be shown by 
marking that is readable from a distance of 30 inches;
    (5) Each passenger entry door operating handle must--
    (i) Be self-illuminated with an initial brightness of at least 160 
microlamberts; or
    (ii) Be conspicuously located and well illuminated by the emergency 
lighting even in conditions of occupant crowding at the door;
    (6) Each passenger entry door with a locking mechanism that is 
released by rotary motion of the handle must be marked--
    (i) With a red arrow, with a shaft of at least three-fourths of an 
inch wide and a head twice the width of the shaft, extending along at 
least 70 degrees of arc at a radius approximately equal to three-fourths 
of the handle length;
    (ii) So that the center line of the exit handle is within one inch of the projected point of the arrow when the 
handle has reached full travel and has released the locking mechanism;
    (iii) With the word ``open'' in red letters, one inch high, placed 
horizontally near the head of the arrow; and
    (7) In addition to the requirements of paragraph (a) of this 
section, the external marking of each emergency exit must--
    (i) Include a 2-inch colorband outlining the exit; and
    (ii) Have a color contrast that is readily distinguishable from the 
surrounding fuselage surface. The contrast must be such that if the 
reflectance of the darker color is 15 percent or less, the reflectance 
of the lighter color must be at least 45 percent. ``Reflectance'' is the 
ratio of the luminous flux reflected by a body to the luminous flux it 
receives. When the reflectance of the darker color is greater than 15 
percent, at least a 30 percent difference between its reflectance and 
the reflectance of the lighter color must be provided.

[Amdt. 23-36, 53 FR 30814, Aug. 15, 1988; 53 FR 34194, Sept. 2, 1988, as 
amended by Amdt. 23-46, 59 FR 25773, May 17, 1994]



Sec. 23.812  Emergency lighting.

    When certification to the emergency exit provisions of Sec. 
23.807(d)(4) is requested, the following apply:
    (a) An emergency lighting system, independent of the main cabin 
lighting system, must be installed. However, the source of general cabin 
illumination may be common to both the emergency and main lighting 
systems if the power supply to the emergency lighting system is 
independent of the power supply to the main lighting system.
    (b) There must be a crew warning light that illuminates in the 
cockpit when power is on in the airplane and the emergency lighting 
control device is not armed.
    (c) The emergency lights must be operable manually from the 
flightcrew station and be provided with automatic activation. The 
cockpit control device must have ``on,'' ``off,'' and ``armed'' 
positions so that, when armed in the cockpit, the lights will operate by 
automatic activation.
    (d) There must be a means to safeguard against inadvertent operation 
of the cockpit control device from the ``armed'' or ``on'' positions.
    (e) The cockpit control device must have provisions to allow the 
emergency lighting system to be armed or activated at any time that it 
may be needed.
    (f) When armed, the emergency lighting system must activate and 
remain lighted when--
    (1) The normal electrical power of the airplane is lost; or
    (2) The airplane is subjected to an impact that results in a 
deceleration in excess of 2g and a velocity change in excess of 3.5 
feet-per-second, acting along the longitudinal axis of the airplane; or
    (3) Any other emergency condition exists where automatic activation 
of the emergency lighting is necessary to aid with occupant evacuation.
    (g) The emergency lighting system must be capable of being turned 
off and reset by the flightcrew after automatic activation.
    (h) The emergency lighting system must provide internal lighting, 
including--

[[Page 260]]

    (1) Illuminated emergency exit marking and locating signs, including 
those required in Sec. 23.811(b);
    (2) Sources of general illumination in the cabin that provide an 
average illumination of not less than 0.05 foot-candle and an 
illumination at any point of not less than 0.01 foot-candle when 
measured along the center line of the main passenger aisle(s) and at the 
seat armrest height; and
    (3) Floor proximity emergency escape path marking that provides 
emergency evacuation guidance for the airplane occupants when all 
sources of illumination more than 4 feet above the cabin aisle floor are 
totally obscured.
    (i) The energy supply to each emergency lighting unit must provide 
the required level of illumination for at least 10 minutes at the 
critical ambient conditions after activation of the emergency lighting 
system.
    (j) If rechargeable batteries are used as the energy supply for the 
emergency lighting system, they may be recharged from the main 
electrical power system of the airplane provided the charging circuit is 
designed to preclude inadvertent battery discharge into the charging 
circuit faults. If the emergency lighting system does not include a 
charging circuit, battery condition monitors are required.
    (k) Components of the emergency lighting system, including 
batteries, wiring, relays, lamps, and switches, must be capable of 
normal operation after being subjected to the inertia forces resulting 
from the ultimate load factors prescribed in Sec. 23.561(b)(2).
    (l) The emergency lighting system must be designed so that after any 
single transverse vertical separation of the fuselage during a crash 
landing:
    (1) At least 75 percent of all electrically illuminated emergency 
lights required by this section remain operative; and
    (2) Each electrically illuminated exit sign required by Sec. 23.811 
(b) and (c) remains operative, except those that are directly damaged by 
the fuselage separation.

[Doc. No. 26324, 59 FR 25774, May 17, 1994]



Sec. 23.813  Emergency exit access.

    (a) For commuter category airplanes, access to window-type emergency 
exits may not be obstructed by seats or seat backs.
    (b) In addition, when certification to the emergency exit provisions 
of Sec. 23.807(d)(4) is requested, the following emergency exit access 
must be provided:
    (1) The passageway leading from the aisle to the passenger entry 
door must be unobstructed and at least 20 inches wide.
    (2) There must be enough space next to the passenger entry door to 
allow assistance in evacuation of passengers without reducing the 
unobstructed width of the passageway below 20 inches.
    (3) If it is necessary to pass through a passageway between 
passenger compartments to reach a required emergency exit from any seat 
in the passenger cabin, the passageway must be unobstructed; however, 
curtains may be used if they allow free entry through the passageway.
    (4) No door may be installed in any partition between passenger 
compartments unless that door has a means to latch it in the open 
position. The latching means must be able to withstand the loads imposed 
upon it by the door when the door is subjected to the inertia loads 
resulting from the ultimate static load factors prescribed in Sec. 
23.561(b)(2).
    (5) If it is necessary to pass through a doorway separating the 
passenger cabin from other areas to reach a required emergency exit from 
any passenger seat, the door must have a means to latch it in the open 
position. The latching means must be able to withstand the loads imposed 
upon it by the door when the door is subjected to the inertia loads 
resulting from the ultimate static load factors prescribed in Sec. 
23.561(b)(2).

[Amdt. 23-36, 53 FR 30815, Aug. 15, 1988, as amended by Amdt. 23-46, 59 
FR 25774, May 17, 1994]



Sec. 23.815  Width of aisle.

    (a) Except as provided in paragraph (b) of this section, for 
commuter category airplanes, the width of the main passenger aisle at 
any point between seats must equal or exceed the values in the following 
table:

[[Page 261]]



------------------------------------------------------------------------
                                    Minimum main passenger aisle width
                                 ---------------------------------------
    Number of passenger seats        Less than 25     25 inches and more
                                   inches from floor      from floor
------------------------------------------------------------------------
10 through 19...................  9 inches..........  15 inches.
------------------------------------------------------------------------

    (b) When certification to the emergency exist provisions of Sec. 
23.807(d)(4) is requested, the main passenger aisle width at any point 
between the seats must equal or exceed the following values:

------------------------------------------------------------------------
                                                 Minimum main passenger
                                                  aisle width (inches)
                                               -------------------------
           Number of passenger seats             Less than    25 inches
                                                 25 inches     and more
                                                 from floor   from floor
------------------------------------------------------------------------
10 or fewer...................................       \1\ 12           15
11 through 19.................................           12           20
------------------------------------------------------------------------
\1\ A narrower width not less than 9 inches may be approved when
  substantiated by tests found necessary by the Administrator.


[Amdt. 23-34, 52 FR 1831, Jan. 15, 1987, as amended by Amdt. 23-46, 59 
FR 25774, May 17, 1994]



Sec. 23.831  Ventilation.

    (a) Each passenger and crew compartment must be suitably ventilated. 
Carbon monoxide concentration may not exceed one part in 20,000 parts of 
air.
    (b) For pressurized airplanes, the ventilating air in the flightcrew 
and passenger compartments must be free of harmful or hazardous 
concentrations of gases and vapors in normal operations and in the event 
of reasonably probable failures or malfunctioning of the ventilating, 
heating, pressurization, or other systems and equipment. If accumulation 
of hazardous quantities of smoke in the cockpit area is reasonably 
probable, smoke evacuation must be readily accomplished starting with 
full pressurization and without depressurizing beyond safe limits.
    (c) For jet pressurized airplanes that operate at altitudes above 
41,000 feet, under normal operating conditions and in the event of any 
probable failure conditions of any system which would adversely affect 
the ventilating air, the ventilation system must provide reasonable 
passenger comfort. The ventilation system must also provide a sufficient 
amount of uncontaminated air to enable the flight crew members to 
perform their duties without undue discomfort or fatigue. For normal 
operating conditions, the ventilation system must be designed to provide 
each occupant with at least 0.55 pounds of fresh air per minute. In the 
event of the loss of one source of fresh air, the supply of fresh 
airflow may not be less than 0.4 pounds per minute for any period 
exceeding five minutes.
    (d) For jet pressurized airplanes that operate at altitudes above 
41,000 feet, other probable and improbable Environmental Control System 
failure conditions that adversely affect the passenger and flight crew 
compartment environmental conditions may not affect flight crew 
performance so as to result in a hazardous condition, and no occupant 
shall sustain permanent physiological harm.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-34, 52 FR 1831, Jan. 15, 1987; Amdt. 23-42, 56 FR 
354, Jan. 3, 1991; Amdt. 23-62, 76 FR 75757, Dec. 2, 2011]

                             Pressurization



Sec. 23.841  Pressurized cabins.

    (a) If certification for operation above 25,000 feet is requested, 
the airplane must be able to maintain a cabin pressure altitude of not 
more than 15,000 feet, in the event of any probable failure condition in 
the pressurization system. During decompression, the cabin altitude may 
not exceed 15,000 feet for more than 10 seconds and 25,000 feet for any 
duration.
    (b) Pressurized cabins must have at least the following valves, 
controls, and indicators, for controlling cabin pressure:
    (1) Two pressure relief valves to automatically limit the positive 
pressure differential to a predetermined value at the maximum rate of 
flow delivered by the pressure source. The combined capacity of the 
relief valves must be large enough so that the failure of any one valve 
would not cause an appreciable rise in the pressure differential. The 
pressure differential is positive when the internal pressure is greater 
than the external.
    (2) Two reverse pressure differential relief valves (or their 
equivalent) to automatically prevent a negative pressure differential 
that would damage the structure. However, one valve is

[[Page 262]]

enough if it is of a design that reasonably precludes its 
malfunctioning.
    (3) A means by which the pressure differential can be rapidly 
equalized.
    (4) An automatic or manual regulator for controlling the intake or 
exhaust airflow, or both, for maintaining the required internal 
pressures and airflow rates.
    (5) Instruments to indicate to the pilot the pressure differential, 
the cabin pressure altitude, and the rate of change of cabin pressure 
altitude.
    (6) Warning indication at the pilot station to indicate when the 
safe or preset pressure differential is exceeded and when a cabin 
pressure altitude of 10,000 feet is exceeded. The 10,000 foot cabin 
altitude warning may be increased up to 15,000 feet for operations from 
high altitude airfields (10,000 to 15,000 feet) provided:
    (i) The landing or the take off modes (normal or high altitude) are 
clearly indicated to the flight crew.
    (ii) Selection of normal or high altitude airfield mode requires no 
more than one flight crew action and goes to normal airfield mode at 
engine stop.
    (iii) The pressurization system is designed to ensure cabin altitude 
does not exceed 10,000 feet when in flight above flight level (FL) 250.
    (iv) The pressurization system and cabin altitude warning system is 
designed to ensure cabin altitude warning at 10,000 feet when in flight 
above FL250.
    (7) A warning placard for the pilot if the structure is not designed 
for pressure differentials up to the maximum relief valve setting in 
combination with landing loads.
    (8) A means to stop rotation of the compressor or to divert airflow 
from the cabin if continued rotation of an engine-driven cabin 
compressor or continued flow of any compressor bleed air will create a 
hazard if a malfunction occurs.
    (c) If certification for operation above 41,000 feet and not more 
than 45,000 feet is requested--
    (1) The airplane must prevent cabin pressure altitude from exceeding 
the following after decompression from any probable pressurization 
system failure in conjunction with any undetected, latent pressurization 
system failure condition:
    (i) If depressurization analysis shows that the cabin altitude does 
not exceed 25,000 feet, the pressurization system must prevent the cabin 
altitude from exceeding the cabin altitude-time history shown in Figure 
1 of this section.
    (ii) Maximum cabin altitude is limited to 30,000 feet. If cabin 
altitude exceeds 25,000 feet, the maximum time the cabin altitude may 
exceed 25,000 feet is 2 minutes; time starting when the cabin altitude 
exceeds 25,000 feet and ending when it returns to 25,000 feet.
    (2) The airplane must prevent cabin pressure altitude from exceeding 
the following after decompression from any single pressurization system 
failure in conjunction with any probable fuselage damage:
    (i) If depressurization analysis shows that the cabin altitude does 
not exceed 37,000 feet, the pressurization system must prevent the cabin 
altitude from exceeding the cabin altitude-time history shown in Figure 
2 of this section.
    (ii) Maximum cabin altitude is limited to 40,000 feet. If cabin 
altitude exceeds 37,000 feet, the maximum time the cabin altitude may 
exceed 25,000 feet is 2 minutes; time starting when the cabin altitude 
exceeds 25,000 feet and ending when it returns to 25,000 feet.
    (3) In showing compliance with paragraphs (c)(1) and (c)(2) of this 
section, it may be assumed that an emergency descent is made by an 
approved emergency procedure. A 17-second flight crew recognition and 
reaction time must be applied between cabin altitude warning and the 
initiation of an emergency descent. Fuselage structure, engine and 
system failures are to be considered in evaluating the cabin 
decompression.

[[Page 263]]

[GRAPHIC] [TIFF OMITTED] TR02DE11.084

[GRAPHIC] [TIFF OMITTED] TR02DE11.085

    (d) If certification for operation above 45,000 feet and not more 
than 51,000 feet is requested--
    (1) Pressurized cabins must be equipped to provide a cabin pressure 
altitude of not more than 8,000 feet at the maximum operating altitude 
of the

[[Page 264]]

airplane under normal operating conditions.
    (2) The airplane must prevent cabin pressure altitude from exceeding 
the following after decompression from any failure condition not shown 
to be extremely improbable:
    (i) Twenty-five thousand (25,000) feet for more than 2 minutes; or
    (ii) Forty thousand (40,000) feet for any duration.
    (3) Fuselage structure, engine and system failures are to be 
considered in evaluating the cabin decompression.
    (4) In addition to the cabin altitude indicating means in (b)(6) of 
this section, an aural or visual signal must be provided to warn the 
flight crew when the cabin pressure altitude exceeds 10,000 feet.
    (5) The sensing system and pressure sensors necessary to meet the 
requirements of (b)(5), (b)(6), and (d)(4) of this section and Sec. 
23.1447(e), must, in the event of low cabin pressure, actuate the 
required warning and automatic presentation devices without any delay 
that would significantly increase the hazards resulting from 
decompression.

[Amdt. 23-14, 38 FR 31822, Nov. 19, 1973, as amended by Amdt. 23-17, 41 
FR 55464, Dec. 20, 1976; Amdt. 23-49, 61 FR 5167, Feb. 9, 1996; Amdt. 
23-62, 76 FR 75757, Dec. 2, 2011]



Sec. 23.843  Pressurization tests.

    (a) Strength test. The complete pressurized cabin, including doors, 
windows, canopy, and valves, must be tested as a pressure vessel for the 
pressure differential specified in Sec. 23.365(d).
    (b) Functional tests. The following functional tests must be 
performed:
    (1) Tests of the functioning and capacity of the positive and 
negative pressure differential valves, and of the emergency release 
valve, to simulate the effects of closed regulator valves.
    (2) Tests of the pressurization system to show proper functioning 
under each possible condition of pressure, temperature, and moisture, up 
to the maximum altitude for which certification is requested.
    (3) Flight tests, to show the performance of the pressure supply, 
pressure and flow regulators, indicators, and warning signals, in steady 
and stepped climbs and descents at rates corresponding to the maximum 
attainable within the operating limitations of the airplane, up to the 
maximum altitude for which certification is requested.
    (4) Tests of each door and emergency exit, to show that they operate 
properly after being subjected to the flight tests prescribed in 
paragraph (b)(3) of this section.

                             Fire Protection



Sec. 23.851  Fire extinguishers.

    (a) There must be at least one hand fire extinguisher for use in the 
pilot compartment that is located within easy access of the pilot while 
seated.
    (b) There must be at least one hand fire extinguisher located 
conveniently in the passenger compartment--
    (1) Of each airplane accommodating more than 6 passengers; and
    (2) Of each commuter category airplane.
    (c) For hand fire extinguishers, the following apply:
    (1) The type and quantity of each extinguishing agent used must be 
appropriate to the kinds of fire likely to occur where that agent is to 
be used.
    (2) Each extinguisher for use in a personnel compartment must be 
designed to minimize the hazard of toxic gas concentrations.

[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]



Sec. 23.853  Passenger and crew compartment interiors.

    For each compartment to be used by the crew or passengers:
    (a) The materials must be at least flame-resistant;
    (b) [Reserved]
    (c) If smoking is to be prohibited, there must be a placard so 
stating, and if smoking is to be allowed--
    (1) There must be an adequate number of self-contained, removable 
ashtrays; and
    (2) Where the crew compartment is separated from the passenger 
compartment, there must be at least one illuminated sign (using either 
letters or symbols) notifying all passengers when smoking is prohibited. 
Signs which notify when smoking is prohibited must--
    (i) When illuminated, be legible to each passenger seated in the 
passenger cabin under all probable lighting conditions; and

[[Page 265]]

    (ii) Be so constructed that the crew can turn the illumination on 
and off; and
    (d) In addition, for commuter category airplanes the following 
requirements apply:
    (1) Each disposal receptacle for towels, paper, or waste must be 
fully enclosed and constructed of at least fire resistant materials and 
must contain fires likely to occur in it under normal use. The ability 
of the disposal receptacle to contain those fires under all probable 
conditions of wear, misalignment, and ventilation expected in service 
must be demonstrated by test. A placard containing the legible words 
``No Cigarette Disposal'' must be located on or near each disposal 
receptacle door.
    (2) Lavatories must have ``No Smoking'' or ``No Smoking in 
Lavatory'' placards located conspicuously on each side of the entry 
door.
    (3) Materials (including finishes or decorative surfaces applied to 
the materials) used in each compartment occupied by the crew or 
passengers must meet the following test criteria as applicable:
    (i) Interior ceiling panels, interior wall panels, partitions, 
galley structure, large cabinet walls, structural flooring, and 
materials used in the construction of stowage compartments (other than 
underseat stowage compartments and compartments for stowing small items 
such as magazines and maps) must be self-extinguishing when tested 
vertically in accordance with the applicable portions of appendix F of 
this part or by other equivalent methods. The average burn length may 
not exceed 6 inches and the average flame time after removal of the 
flame source may not exceed 15 seconds. Drippings from the test specimen 
may not continue to flame for more than an average of 3 seconds after 
falling.
    (ii) Floor covering, textiles (including draperies and upholstery), 
seat cushions, padding, decorative and nondecorative coated fabrics, 
leather, trays and galley furnishings, electrical conduit, thermal and 
acoustical insulation and insulation covering, air ducting, joint and 
edge covering, cargo compartment liners, insulation blankets, cargo 
covers and transparencies, molded and thermoformed parts, air ducting 
joints, and trim strips (decorative and chafing), that are constructed 
of materials not covered in paragraph (d)(3)(iv) of this section must be 
self extinguishing when tested vertically in accordance with the 
applicable portions of appendix F of this part or other approved 
equivalent methods. The average burn length may not exceed 8 inches and 
the average flame time after removal of the flame source may not exceed 
15 seconds. Drippings from the test specimen may not continue to flame 
for more than an average of 5 seconds after falling.
    (iii) Motion picture film must be safety film meeting the Standard 
Specifications for Safety Photographic Film PH1.25 (available from the 
American National Standards Institute, 1430 Broadway, New York, N.Y. 
10018) or an FAA approved equivalent. If the film travels through ducts, 
the ducts must meet the requirements of paragraph (d)(3)(ii) of this 
section.
    (iv) Acrylic windows and signs, parts constructed in whole or in 
part of elastomeric materials, edge-lighted instrument assemblies 
consisting of two or more instruments in a common housing, seatbelts, 
shoulder harnesses, and cargo and baggage tiedown equipment, including 
containers, bins, pallets, etc., used in passenger or crew compartments, 
may not have an average burn rate greater than 2.5 inches per minute 
when tested horizontally in accordance with the applicable portions of 
appendix F of this part or by other approved equivalent methods.
    (v) Except for electrical wire cable insulation, and for small parts 
(such as knobs, handles, rollers, fasteners, clips, grommets, rub 
strips, pulleys, and small electrical parts) that the Administrator 
finds would not contribute significantly to the propagation of a fire, 
materials in items not specified in paragraphs (d)(3)(i), (ii), (iii), 
or (iv) of this section may not have a burn rate greater than 4.0 inches 
per minute when tested horizontally in accordance with the applicable 
portions of appendix F of this part or by other approved equivalent 
methods.
    (e) Lines, tanks, or equipment containing fuel, oil, or other 
flammable fluids may not be installed in such

[[Page 266]]

compartments unless adequately shielded, isolated, or otherwise 
protected so that any breakage or failure of such an item would not 
create a hazard.
    (f) Airplane materials located on the cabin side of the firewall 
must be self-extinguishing or be located at such a distance from the 
firewall, or otherwise protected, so that ignition will not occur if the 
firewall is subjected to a flame temperature of not less than 2,000 
degrees F for 15 minutes. For self-extinguishing materials (except 
electrical wire and cable insulation and small parts that the 
Administrator finds would not contribute significantly to the 
propagation of a fire), a vertifical self-extinguishing test must be 
conducted in accordance with appendix F of this part or an equivalent 
method approved by the Administrator. The average burn length of the 
material may not exceed 6 inches and the average flame time after 
removal of the flame source may not exceed 15 seconds. Drippings from 
the material test specimen may not continue to flame for more than an 
average of 3 seconds after falling.

[Amdt. 23-14, 23 FR 31822, Nov. 19, 1973, as amended by Amdt. 23-23, 43 
FR 50593, Oct. 30, 1978; Amdt. 23-25, 45 FR 7755, Feb. 4, 1980; Amdt. 
23-34, 52 FR 1831, Jan. 15, 1987; Amdt. 23-62, 76 FR 75759, Dec. 2, 
2011]



Sec. 23.855  Cargo and baggage compartment fire protection.

    (a) Sources of heat within each cargo and baggage compartment that 
are capable of igniting the compartment contents must be shielded and 
insulated to prevent such ignition.
    (b) Each cargo and baggage compartment must be constructed of 
materials that meet the appropriate provisions of Sec. 23.853(d)(3).
    (c) In addition, for commuter category airplanes, each cargo and 
baggage compartment must:
    (1) Be located where the presence of a fire would be easily 
discovered by the pilots when seated at their duty station, or it must 
be equipped with a smoke or fire detector system to give a warning at 
the pilots' station, and provide sufficient access to enable a pilot to 
effectively reach any part of the compartment with the contents of a 
hand held fire extinguisher, or
    (2) Be equipped with a smoke or fire detector system to give a 
warning at the pilots' station and have ceiling and sidewall liners and 
floor panels constructed of materials that have been subjected to and 
meet the 45 degree angle test of appendix F of this part. The flame may 
not penetrate (pass through) the material during application of the 
flame or subsequent to its removal. The average flame time after removal 
of the flame source may not exceed 15 seconds, and the average glow time 
may not exceed 10 seconds. The compartment must be constructed to 
provide fire protection that is not less than that required of its 
individual panels; or
    (3) Be constructed and sealed to contain any fire within the 
compartment.

[Doc. No. 27806, 61 FR 5167, Feb. 9, 1996]



Sec. 23.856  Thermal/acoustic insulation materials.

    Thermal/acoustic insulation material installed in the fuselage must 
meet the flame propagation test requirements of part II of Appendix F to 
this part, or other approved equivalent test requirements. This 
requirement does not apply to ``small parts,'' as defined in Sec. 
23.853(d)(3)(v).

[Amdt. 23-62, 76 FR 75759, Dec. 2, 2011]



Sec. 23.859  Combustion heater fire protection.

    (a) Combustion heater fire regions. The following combustion heater 
fire regions must be protected from fire in accordance with the 
applicable provisions of Sec. Sec. 23.1182 through 23.1191 and 23.1203:
    (1) The region surrounding the heater, if this region contains any 
flammable fluid system components (excluding the heater fuel system) 
that could--
    (i) Be damaged by heater malfunctioning; or
    (ii) Allow flammable fluids or vapors to reach the heater in case of 
leakage.
    (2) The region surrounding the heater, if the heater fuel system has 
fittings that, if they leaked, would allow fuel vapor to enter this 
region.

[[Page 267]]

    (3) The part of the ventilating air passage that surrounds the 
combustion chamber.
    (b) Ventilating air ducts. Each ventilating air duct passing through 
any fire region must be fireproof. In addition--
    (1) Unless isolation is provided by fireproof valves or by equally 
effective means, the ventilating air duct downstream of each heater must 
be fireproof for a distance great enough to ensure that any fire 
originating in the heater can be contained in the duct; and
    (2) Each part of any ventilating duct passing through any region 
having a flammable fluid system must be constructed or isolated from 
that system so that the malfunctioning of any component of that system 
cannot introduce flammable fluids or vapors into the ventilating 
airstream.
    (c) Combustion air ducts. Each combustion air duct must be fireproof 
for a distance great enough to prevent damage from backfiring or reverse 
flame propagation. In addition--
    (1) No combustion air duct may have a common opening with the 
ventilating airstream unless flames from backfires or reverse burning 
cannot enter the ventilating airstream under any operating condition, 
including reverse flow or malfunctioning of the heater or its associated 
components; and
    (2) No combustion air duct may restrict the prompt relief of any 
backfire that, if so restricted, could cause heater failure.
    (d) Heater controls: general. Provision must be made to prevent the 
hazardous accumulation of water or ice on or in any heater control 
component, control system tubing, or safety control.
    (e) Heater safety controls. (1) Each combustion heater must have the 
following safety controls:
    (i) Means independent of the components for the normal continuous 
control of air temperature, airflow, and fuel flow must be provided to 
automatically shut off the ignition and fuel supply to that heater at a 
point remote from that heater when any of the following occurs:
    (A) The heater exchanger temperature exceeds safe limits.
    (B) The ventilating air temperature exceeds safe limits.
    (C) The combustion airflow becomes inadequate for safe operation.
    (D) The ventilating airflow becomes inadequate for safe operation.
    (ii) Means to warn the crew when any heater whose heat output is 
essential for safe operation has been shut off by the automatic means 
prescribed in paragraph (e)(1)(i) of this section.
    (2) The means for complying with paragraph (e)(1)(i) of this section 
for any individual heater must--
    (i) Be independent of components serving any other heater whose heat 
output is essential for safe operations; and
    (ii) Keep the heater off until restarted by the crew.
    (f) Air intakes. Each combustion and ventilating air intake must be 
located so that no flammable fluids or vapors can enter the heater 
system under any operating condition--
    (1) During normal operation; or
    (2) As a result of the malfunctioning of any other component.
    (g) Heater exhaust. Heater exhaust systems must meet the provisions 
of Sec. Sec. 23.1121 and 23.1123. In addition, there must be provisions 
in the design of the heater exhaust system to safely expel the products 
of combustion to prevent the occurrence of--
    (1) Fuel leakage from the exhaust to surrounding compartments;
    (2) Exhaust gas impingement on surrounding equipment or structure;
    (3) Ignition of flammable fluids by the exhaust, if the exhaust is 
in a compartment containing flammable fluid lines; and
    (4) Restrictions in the exhaust system to relieve backfires that, if 
so restricted, could cause heater failure.
    (h) Heater fuel systems. Each heater fuel system must meet each 
powerplant fuel system requirement affecting safe heater operation. Each 
heater fuel system component within the ventilating airstream must be 
protected by shrouds so that no leakage from those components can enter 
the ventilating airstream.
    (i) Drains. There must be means to safely drain fuel that might 
accumulate within the combustion chamber or the heater exchanger. In 
addition--

[[Page 268]]

    (1) Each part of any drain that operates at high temperatures must 
be protected in the same manner as heater exhausts; and
    (2) Each drain must be protected from hazardous ice accumulation 
under any operating condition.

[Amdt. 23-27, 45 FR 70387, Oct. 23, 1980]



Sec. 23.863  Flammable fluid fire protection.

    (a) In each area where flammable fluids or vapors might escape by 
leakage of a fluid system, there must be means to minimize the 
probability of ignition of the fluids and vapors, and the resultant 
hazard if ignition does occur.
    (b) Compliance with paragraph (a) of this section must be shown by 
analysis or tests, and the following factors must be considered:
    (1) Possible sources and paths of fluid leakage, and means of 
detecting leakage.
    (2) Flammability characteristics of fluids, including effects of any 
combustible or absorbing materials.
    (3) Possible ignition sources, including electrical faults, 
overheating of equipment, and malfunctioning of protective devices.
    (4) Means available for controlling or extinguishing a fire, such as 
stopping flow of fluids, shutting down equipment, fireproof containment, 
or use of extinguishing agents.
    (5) Ability of airplane components that are critical to safety of 
flight to withstand fire and heat.
    (c) If action by the flight crew is required to prevent or 
counteract a fluid fire (e.g. equipment shutdown or actuation of a fire 
extinguisher), quick acting means must be provided to alert the crew.
    (d) Each area where flammable fluids or vapors might escape by 
leakage of a fluid system must be identified and defined.

[Amdt. 23-23, 43 FR 50593, Oct. 30, 1978]



Sec. 23.865  Fire protection of flight controls, engine mounts,
and other flight structure.

    Flight controls, engine mounts, and other flight structure located 
in designated fire zones, or in adjacent areas that would be subjected 
to the effects of fire in the designated fire zones, must be constructed 
of fireproof material or be shielded so that they are capable of 
withstanding the effects of a fire. Engine vibration isolators must 
incorporate suitable features to ensure that the engine is retained if 
the non-fireproof portions of the isolators deteriorate from the effects 
of a fire.

[Doc. No. 27805, 61 FR 5148, Feb. 9, 1996]

               Electrical Bonding and Lightning Protection



Sec. 23.867  Electrical bonding and protection against lightning and
static electricity.

    (a) The airplane must be protected against catastrophic effects from 
lightning.
    (b) For metallic components, compliance with paragraph (a) of this 
section may be shown by--
    (1) Bonding the components properly to the airframe; or
    (2) Designing the components so that a strike will not endanger the 
airplane.
    (c) For nonmetallic components, compliance with paragraph (a) of 
this section may be shown by--
    (1) Designing the components to minimize the effect of a strike; or
    (2) Incorporating acceptable means of diverting the resulting 
electrical current so as not to endanger the airplane.

[Amdt. 23-7, 34 FR 13092, Aug. 13, 1969]

                              Miscellaneous



Sec. 23.871  Leveling means.

    There must be means for determining when the airplane is in a level 
position on the ground.

[Amdt. 23-7, 34 FR 13092, Aug. 13, 1969]



                          Subpart E_Powerplant

                                 General



Sec. 23.901  Installation.

    (a) For the purpose of this part, the airplane powerplant 
installation includes each component that--
    (1) Is necessary for propulsion; and
    (2) Affects the safety of the major propulsive units.
    (b) Each powerplant installation must be constructed and arranged 
to--

[[Page 269]]

    (1) Ensure safe operation to the maximum altitude for which approval 
is requested.
    (2) Be accessible for necessary inspections and maintenance.
    (c) Engine cowls and nacelles must be easily removable or openable 
by the pilot to provide adequate access to and exposure of the engine 
compartment for preflight checks.
    (d) Each turbine engine installation must be constructed and 
arranged to--
    (1) Result in carcass vibration characteristics that do not exceed 
those established during the type certification of the engine.
    (2) Ensure that the capability of the installed engine to withstand 
the ingestion of rain, hail, ice, and birds into the engine inlet is not 
less than the capability established for the engine itself under Sec. 
23.903(a)(2).
    (e) The installation must comply with--
    (1) The instructions provided under the engine type certificate and 
the propeller type certificate.
    (2) The applicable provisions of this subpart.
    (f) Each auxiliary power unit installation must meet the applicable 
portions of this part.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13092, Aug. 13, 1969; Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 
23-29, 49 FR 6846, Feb. 23, 1984; Amdt. 23-34, 52 FR 1832, Jan. 15, 
1987; Amdt. 23-34, 52 FR 34745, Sept. 14, 1987; Amdt. 23-43, 58 FR 
18970, Apr. 9, 1993; Amdt. 23-51, 61 FR 5136, Feb. 9, 1996; Amdt. 23-53, 
63 FR 14797, Mar. 26, 1998]



Sec. 23.903  Engines.

    (a) Engine type certificate. (1) Each engine must have a type 
certificate and must meet the applicable requirements of part 34 of this 
chapter.
    (2) Each turbine engine and its installation must comply with one of 
the following:
    (i) Sections 33.76, 33.77 and 33.78 of this chapter in effect on 
December 13, 2000, or as subsequently amended; or
    (ii) Sections 33.77 and 33.78 of this chapter in effect on April 30, 
1998, or as subsequently amended before December 13, 2000; or
    (iii) Section 33.77 of this chapter in effect on October 31, 1974, 
or as subsequently amended before April 30, 1998, unless that engine's 
foreign object ingestion service history has resulted in an unsafe 
condition; or
    (iv) Be shown to have a foreign object ingestion service history in 
similar installation locations which has not resulted in any unsafe 
condition.

    Note: Sec. 33.77 of this chapter in effect on October 31, 1974, was 
published in 14 CFR parts 1 to 59, Revised as of January 1, 1975. See 39 
FR 35467, October 1, 1974.

    (b) Turbine engine installations. For turbine engine installations--
    (1) Design precautions must be taken to minimize the hazards to the 
airplane in the event of an engine rotor failure or of a fire 
originating inside the engine which burns through the engine case.
    (2) The powerplant systems associated with engine control devices, 
systems, and instrumentation must be designed to give reasonable 
assurance that those operating limitations that adversely affect turbine 
rotor structural integrity will not be exceeded in service.
    (3) For engines embedded in the fuselage behind the cabin, the 
effects of a fan exiting forward of the inlet case (fan disconnect) must 
be addressed, the passengers must be protected, and the airplane must be 
controllable to allow for continued safe flight and landing.
    (c) Engine isolation. The powerplants must be arranged and isolated 
from each other to allow operation, in at least one configuration, so 
that the failure or malfunction of any engine, or the failure or 
malfunction (including destruction by fire in the engine compartment) of 
any system that can affect an engine (other than a fuel tank if only one 
fuel tank is installed), will not:
    (1) Prevent the continued safe operation of the remaining engines; 
or
    (2) Require immediate action by any crewmember for continued safe 
operation of the remaining engines.
    (d) Starting and stopping (piston engine). (1) The design of the 
installation must be such that risk of fire or mechanical damage to the 
engine or airplane, as a result of starting the engine in any conditions 
in which starting is to be permitted, is reduced to a minimum. Any 
techniques and associated limitations for engine starting must be

[[Page 270]]

established and included in the Airplane Flight Manual, approved manual 
material, or applicable operating placards. Means must be provided for--
    (i) Restarting any engine of a multiengine airplane in flight, and
    (ii) Stopping any engine in flight, after engine failure, if 
continued engine rotation would cause a hazard to the airplane.
    (2) In addition, for commuter category airplanes, the following 
apply:
    (i) Each component of the stopping system on the engine side of the 
firewall that might be exposed to fire must be at least fire resistant.
    (ii) If hydraulic propeller feathering systems are used for this 
purpose, the feathering lines must be at least fire resistant under the 
operating conditions that may be expected to exist during feathering.
    (e) Starting and stopping (turbine engine). Turbine engine 
installations must comply with the following:
    (1) The design of the installation must be such that risk of fire or 
mechanical damage to the engine or the airplane, as a result of starting 
the engine in any conditions in which starting is to be permitted, is 
reduced to a minimum. Any techniques and associated limitations must be 
established and included in the Airplane Flight Manual, approved manual 
material, or applicable operating placards.
    (2) There must be means for stopping combustion within any engine 
and for stopping the rotation of any engine if continued rotation would 
cause a hazard to the airplane. Each component of the engine stopping 
system located in any fire zone must be fire resistant. If hydraulic 
propeller feathering systems are used for stopping the engine, the 
hydraulic feathering lines or hoses must be fire resistant.
    (3) It must be possible to restart an engine in flight. Any 
techniques and associated limitations must be established and included 
in the Airplane Flight Manual, approved manual material, or applicable 
operating placards.
    (4) It must be demonstrated in flight that when restarting engines 
following a false start, all fuel or vapor is discharged in such a way 
that it does not constitute a fire hazard.
    (f) Restart envelope. An altitude and airspeed envelope must be 
established for the airplane for in-flight engine restarting and each 
installed engine must have a restart capability within that envelope.
    (g) Restart capability. For turbine engine powered airplanes, if the 
minimum windmilling speed of the engines, following the in-flight 
shutdown of all engines, is insufficient to provide the necessary 
electrical power for engine ignition, a power source independent of the 
engine-driven electrical power generating system must be provided to 
permit in-flight engine ignition for restarting.

[Amdt. 23-14, 38 FR 31822, Nov. 19, 1973]

    Editorial Note: For Federal Register citations affecting Sec. 
23.903, see the List of CFR Sections Affected, which appears in the 
Finding Aids section of the printed volume and at www.fdsys.gov.



Sec. 23.904  Automatic power reserve system.

    If installed, an automatic power reserve (APR) system that 
automatically advances the power or thrust on the operating engine(s), 
when any engine fails during takeoff, must comply with appendix H of 
this part.

[Doc. No. 26344, 58 FR 18970, Apr. 9, 1993]



Sec. 23.905  Propellers.

    (a) Each propeller must have a type certificate.
    (b) Engine power and propeller shaft rotational speed may not exceed 
the limits for which the propeller is certificated.
    (c) Each featherable propeller must have a means to unfeather it in 
flight.
    (d) The propeller blade pitch control system must meet the 
requirements of Sec. Sec. 35.21, 35.23, 35.42 and 35.43 of this 
chapter.
    (e) All areas of the airplane forward of the pusher propeller that 
are likely to accumulate and shed ice into the propeller disc during any 
operating condition must be suitably protected to prevent ice formation, 
or it must be shown that any ice shed into the propeller disc will not 
create a hazardous condition.
    (f) Each pusher propeller must be marked so that the disc is 
conspicuous

[[Page 271]]

under normal daylight ground conditions.
    (g) If the engine exhaust gases are discharged into the pusher 
propeller disc, it must be shown by tests, or analysis supported by 
tests, that the propeller is capable of continuous safe operation.
    (h) All engine cowling, access doors, and other removable items must 
be designed to ensure that they will not separate from the airplane and 
contact the pusher propeller.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-26, 
45 FR 60171, Sept. 11, 1980; Amdt. 23-29, 49 FR 6847, Feb. 23, 1984; 
Amdt. 23-43, 58 FR 18970, Apr. 9, 1993; Amdt. 23-59, 73 FR 63345, Oct. 
24, 2008]



Sec. 23.907  Propeller vibration and fatigue.

    This section does not apply to fixed-pitch wood propellers of 
conventional design.
    (a) The applicant must determine the magnitude of the propeller 
vibration stresses or loads, including any stress peaks and resonant 
conditions, throughout the operational envelope of the airplane by 
either:
    (1) Measurement of stresses or loads through direct testing or 
analysis based on direct testing of the propeller on the airplane and 
engine installation for which approval is sought; or
    (2) Comparison of the propeller to similar propellers installed on 
similar airplane installations for which these measurements have been 
made.
    (b) The applicant must demonstrate by tests, analysis based on 
tests, or previous experience on similar designs that the propeller does 
not experience harmful effects of flutter throughout the operational 
envelope of the airplane.
    (c) The applicant must perform an evaluation of the propeller to 
show that failure due to fatigue will be avoided throughout the 
operational life of the propeller using the fatigue and structural data 
obtained in accordance with part 35 of this chapter and the vibration 
data obtained from compliance with paragraph (a) of this section. For 
the purpose of this paragraph, the propeller includes the hub, blades, 
blade retention component and any other propeller component whose 
failure due to fatigue could be catastrophic to the airplane. This 
evaluation must include:
    (1) The intended loading spectra including all reasonably 
foreseeable propeller vibration and cyclic load patterns, identified 
emergency conditions, allowable overspeeds and overtorques, and the 
effects of temperatures and humidity expected in service.
    (2) The effects of airplane and propeller operating and 
airworthiness limitations.

[Amdt. 23-59, 73 FR 63345, Oct. 24, 2008]



Sec. 23.909  Turbocharger systems.

    (a) Each turbocharger must be approved under the engine type 
certificate or it must be shown that the turbocharger system, while in 
its normal engine installation and operating in the engine environment--
    (1) Can withstand, without defect, an endurance test of 150 hours 
that meets the applicable requirements of Sec. 33.49 of this 
subchapter; and
    (2) Will have no adverse effect upon the engine.
    (b) Control system malfunctions, vibrations, and abnormal speeds and 
temperatures expected in service may not damage the turbocharger 
compressor or turbine.
    (c) Each turbocharger case must be able to contain fragments of a 
compressor or turbine that fails at the highest speed that is obtainable 
with normal speed control devices inoperative.
    (d) Each intercooler installation, where provided, must comply with 
the following--
    (1) The mounting provisions of the intercooler must be designed to 
withstand the loads imposed on the system;
    (2) It must be shown that, under the installed vibration 
environment, the intercooler will not fail in a manner allowing portions 
of the intercooler to be ingested by the engine; and
    (3) Airflow through the intercooler must not discharge directly on 
any airplane component (e.g., windshield) unless such discharge is shown 
to cause no hazard to the airplane under all operating conditions.
    (e) Engine power, cooling characteristics, operating limits, and 
procedures

[[Page 272]]

affected by the turbocharger system installations must be evaluated. 
Turbocharger operating procedures and limitations must be included in 
the Airplane Flight Manual in accordance with Sec. 23.1581.

[Amdt. 23-7, 34 FR 13092, Aug. 13, 1969, as amended by Amdt. 23-43, 58 
FR 18970, Apr. 9, 1993]



Sec. 23.925  Propeller clearance.

    Unless smaller clearances are substantiated, propeller clearances, 
with the airplane at the most adverse combination of weight and center 
of gravity, and with the propeller in the most adverse pitch position, 
may not be less than the following:
    (a) Ground clearance. There must be a clearance of at least seven 
inches (for each airplane with nose wheel landing gear) or nine inches 
(for each airplane with tail wheel landing gear) between each propeller 
and the ground with the landing gear statically deflected and in the 
level, normal takeoff, or taxing attitude, whichever is most critical. 
In addition, for each airplane with conventional landing gear struts 
using fluid or mechanical means for absorbing landing shocks, there must 
be positive clearance between the propeller and the ground in the level 
takeoff attitude with the critical tire completely deflated and the 
corresponding landing gear strut bottomed. Positive clearance for 
airplanes using leaf spring struts is shown with a deflection 
corresponding to 1.5g.
    (b) Aft-mounted propellers. In addition to the clearances specified 
in paragraph (a) of this section, an airplane with an aft mounted 
propeller must be designed such that the propeller will not contact the 
runway surface when the airplane is in the maximum pitch attitude 
attainable during normal takeoffs and landings.
    (c) Water clearance. There must be a clearance of at least 18 inches 
between each propeller and the water, unless compliance with Sec. 
23.239 can be shown with a lesser clearance.
    (d) Structural clearance. There must be--
    (1) At least one inch radial clearance between the blade tips and 
the airplane structure, plus any additional radial clearance necessary 
to prevent harmful vibration;
    (2) At least one-half inch longitudinal clearance between the 
propeller blades or cuffs and stationary parts of the airplane; and
    (3) Positive clearance between other rotating parts of the propeller 
or spinner and stationary parts of the airplane.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18971, Apr. 9, 1993; Amdt. 23-51, 61 FR 5136, Feb. 9, 1996; Amdt. 
23-48, 61 FR 5148, Feb. 9, 1996]



Sec. 23.929  Engine installation ice protection.

    Propellers (except wooden propellers) and other components of 
complete engine installations must be protected against the accumulation 
of ice as necessary to enable satisfactory functioning without 
appreciable loss of thrust when operated in the icing conditions for 
which certification is requested.

[Amdt. 23-14, 33 FR 31822, Nov. 19, 1973, as amended by Amdt. 23-51, 61 
FR 5136, Feb. 9, 1996]



Sec. 23.933  Reversing systems.

    (a) For turbojet and turbofan reversing systems. (1) Each system 
intended for ground operation only must be designed so that, during any 
reversal in flight, the engine will produce no more than flight idle 
thrust. In addition, it must be shown by analysis or test, or both, 
that--
    (i) Each operable reverser can be restored to the forward thrust 
position; or
    (ii) The airplane is capable of continued safe flight and landing 
under any possible position of the thrust reverser.
    (2) Each system intended for in-flight use must be designed so that 
no unsafe condition will result during normal operation of the system, 
or from any failure, or likely combination of failures, of the reversing 
system under any operating condition including ground operation. Failure 
of structural elements need not be considered if the probability of this 
type of failure is extremely remote.
    (3) Each system must have a means to prevent the engine from 
producing

[[Page 273]]

more than idle thrust when the reversing system malfunctions; except 
that it may produce any greater thrust that is shown to allow 
directional control to be maintained, with aerodynamic means alone, 
under the most critical reversing condition expected in operation.
    (b) For propeller reversing systems. (1) Each system must be 
designed so that no single failure, likely combination of failures or 
malfunction of the system will result in unwanted reverse thrust under 
any operating condition. Failure of structural elements need not be 
considered if the probability of this type of failure is extremely 
remote.
    (2) Compliance with paragraph (b)(1) of this section must be shown 
by failure analysis, or testing, or both, for propeller systems that 
allow the propeller blades to move from the flight low-pitch position to 
a position that is substantially less than the normal flight, low-pitch 
position. The analysis may include or be supported by the analysis made 
to show compliance with Sec. 35.21 for the type certification of the 
propeller and associated installation components. Credit will be given 
for pertinent analysis and testing completed by the engine and propeller 
manufacturers.

[Doc. No. 26344, 58 FR 18971, Apr. 9, 1993, as amended by Amdt. 23-51, 
61 FR 5136, Feb. 9, 1996]



Sec. 23.934  Turbojet and turbofan engine thrust reverser systems tests.

    Thrust reverser systems of turbojet or turbofan engines must meet 
the requirements of Sec. 33.97 of this chapter or it must be 
demonstrated by tests that engine operation and vibratory levels are not 
affected.

[Doc. No. 26344, 58 FR 18971, Apr. 9, 1993]



Sec. 23.937  Turbopropeller-drag limiting systems.

    (a) Turbopropeller-powered airplane propeller-drag limiting systems 
must be designed so that no single failure or malfunction of any of the 
systems during normal or emergency operation results in propeller drag 
in excess of that for which the airplane was designed under the 
structural requirements of this part. Failure of structural elements of 
the drag limiting systems need not be considered if the probability of 
this kind of failure is extremely remote.
    (b) As used in this section, drag limiting systems include manual or 
automatic devices that, when actuated after engine power loss, can move 
the propeller blades toward the feather position to reduce windmilling 
drag to a safe level.

[Amdt. 23-7, 34 FR 13093, Aug. 13, 1969, as amended by Amdt. 23-43, 58 
FR 18971, Apr. 9, 1993]



Sec. 23.939  Powerplant operating characteristics.

    (a) Turbine engine powerplant operating characteristics must be 
investigated in flight to determine that no adverse characteristics 
(such as stall, surge, or flameout) are present, to a hazardous degree, 
during normal and emergency operation within the range of operating 
limitations of the airplane and of the engine.
    (b) Turbocharged reciprocating engine operating characteristics must 
be investigated in flight to assure that no adverse characteristics, as 
a result of an inadvertent overboost, surge, flooding, or vapor lock, 
are present during normal or emergency operation of the engine(s) 
throughout the range of operating limitations of both airplane and 
engine.
    (c) For turbine engines, the air inlet system must not, as a result 
of airflow distortion during normal operation, cause vibration harmful 
to the engine.

[Amdt. 23-7, 34 FR 13093 Aug. 13, 1969, as amended by Amdt. 23-14, 38 FR 
31823, Nov. 19, 1973; Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-
42, 56 FR 354, Jan. 3, 1991]



Sec. 23.943  Negative acceleration.

    No hazardous malfunction of an engine, an auxiliary power unit 
approved for use in flight, or any component or system associated with 
the powerplant or auxiliary power unit may occur when the airplane is 
operated at the negative accelerations within the flight envelopes 
prescribed in Sec. 23.333. This must be shown for the greatest value 
and duration of the acceleration expected in service.

[Amdt. 23-18, 42 FR 15041, Mar. 17, 1977, as amended by Amdt. 23-43, 58 
FR 18971, Apr. 9, 1993]

[[Page 274]]

                               Fuel System



Sec. 23.951  General.

    (a) Each fuel system must be constructed and arranged to ensure fuel 
flow at a rate and pressure established for proper engine and auxiliary 
power unit functioning under each likely operating condition, including 
any maneuver for which certification is requested and during which the 
engine or auxiliary power unit is permitted to be in operation.
    (b) Each fuel system must be arranged so that--
    (1) No fuel pump can draw fuel from more than one tank at a time; or
    (2) There are means to prevent introducing air into the system.
    (c) Each fuel system for a turbine engine must be capable of 
sustained operation throughout its flow and pressure range with fuel 
initially saturated with water at 80 [deg]F and having 0.75cc of free 
water per gallon added and cooled to the most critical condition for 
icing likely to be encountered in operation.
    (d) Each fuel system for a turbine engine powered airplane must meet 
the applicable fuel venting requirements of part 34 of this chapter.

[Amdt. 23-15, 39 FR 35459, Oct. 1, 1974, as amended by Amdt. 23-40, 55 
FR 32861, Aug. 10, 1990; Amdt. 23-43, 58 FR 18971, Apr. 9, 1993]



Sec. 23.953  Fuel system independence.

    (a) Each fuel system for a multiengine airplane must be arranged so 
that, in at least one system configuration, the failure of any one 
component (other than a fuel tank) will not result in the loss of power 
of more than one engine or require immediate action by the pilot to 
prevent the loss of power of more than one engine.
    (b) If a single fuel tank (or series of fuel tanks interconnected to 
function as a single fuel tank) is used on a multiengine airplane, the 
following must be provided:
    (1) Independent tank outlets for each engine, each incorporating a 
shut-off valve at the tank. This shutoff valve may also serve as the 
fire wall shutoff valve required if the line between the valve and the 
engine compartment does not contain more than one quart of fuel (or any 
greater amount shown to be safe) that can escape into the engine 
compartment.
    (2) At least two vents arranged to minimize the probability of both 
vents becoming obstructed simultaneously.
    (3) Filler caps designed to minimize the probability of incorrect 
installation or inflight loss.
    (4) A fuel system in which those parts of the system from each tank 
outlet to any engine are independent of each part of the system 
supplying fuel to any other engine.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13093 Aug. 13, 1969; Amdt. 23-43, 58 FR 18971, Apr. 9, 1993]



Sec. 23.954  Fuel system lightning protection.

    The fuel system must be designed and arranged to prevent the 
ignition of fuel vapor within the system by--
    (a) Direct lightning strikes to areas having a high probability of 
stroke attachment;
    (b) Swept lightning strokes on areas where swept strokes are highly 
probable; and
    (c) Corona or streamering at fuel vent outlets.

[Amdt. 23-7, 34 FR 13093, Aug. 13, 1969]



Sec. 23.955  Fuel flow.

    (a) General. The ability of the fuel system to provide fuel at the 
rates specified in this section and at a pressure sufficient for proper 
engine operation must be shown in the attitude that is most critical 
with respect to fuel feed and quantity of unusable fuel. These 
conditions may be simulated in a suitable mockup. In addition--
    (1) The quantity of fuel in the tank may not exceed the amount 
established as the unusable fuel supply for that tank under Sec. 
23.959(a) plus that quantity necessary to show compliance with this 
section.
    (2) If there is a fuel flowmeter, it must be blocked during the flow 
test and the fuel must flow through the meter or its bypass.
    (3) If there is a flowmeter without a bypass, it must not have any 
probable failure mode that would restrict fuel flow below the level 
required for this fuel demonstration.
    (4) The fuel flow must include that flow necessary for vapor return 
flow,

[[Page 275]]

jet pump drive flow, and for all other purposes for which fuel is used.
    (b) Gravity systems. The fuel flow rate for gravity systems (main 
and reserve supply) must be 150 percent of the takeoff fuel consumption 
of the engine.
    (c) Pump systems. The fuel flow rate for each pump system (main and 
reserve supply) for each reciprocating engine must be 125 percent of the 
fuel flow required by the engine at the maximum takeoff power approved 
under this part.
    (1) This flow rate is required for each main pump and each emergency 
pump, and must be available when the pump is operating as it would 
during takeoff;
    (2) For each hand-operated pump, this rate must occur at not more 
than 60 complete cycles (120 single strokes) per minute.
    (3) The fuel pressure, with main and emergency pumps operating 
simultaneously, must not exceed the fuel inlet pressure limits of the 
engine unless it can be shown that no adverse effect occurs.
    (d) Auxiliary fuel systems and fuel transfer systems. Paragraphs 
(b), (c), and (f) of this section apply to each auxiliary and transfer 
system, except that--
    (1) The required fuel flow rate must be established upon the basis 
of maximum continuous power and engine rotational speed, instead of 
takeoff power and fuel consumption; and
    (2) If there is a placard providing operating instructions, a lesser 
flow rate may be used for transferring fuel from any auxiliary tank into 
a larger main tank. This lesser flow rate must be adequate to maintain 
engine maximum continuous power but the flow rate must not overfill the 
main tank at lower engine powers.
    (e) Multiple fuel tanks. For reciprocating engines that are supplied 
with fuel from more than one tank, if engine power loss becomes apparent 
due to fuel depletion from the tank selected, it must be possible after 
switching to any full tank, in level flight, to obtain 75 percent 
maximum continuous power on that engine in not more than--
    (1) 10 seconds for naturally aspirated single-engine airplanes;
    (2) 20 seconds for turbocharged single-engine airplanes, provided 
that 75 percent maximum continuous naturally aspirated power is regained 
within 10 seconds; or
    (3) 20 seconds for multiengine airplanes.
    (f) Turbine engine fuel systems. Each turbine engine fuel system 
must provide at least 100 percent of the fuel flow required by the 
engine under each intended operation condition and maneuver. The 
conditions may be simulated in a suitable mockup. This flow must--
    (1) Be shown with the airplane in the most adverse fuel feed 
condition (with respect to altitudes, attitudes, and other conditions) 
that is expected in operation; and
    (2) For multiengine airplanes, notwithstanding the lower flow rate 
allowed by paragraph (d) of this section, be automatically uninterrupted 
with respect to any engine until all the fuel scheduled for use by that 
engine has been consumed. In addition--
    (i) For the purposes of this section, ``fuel scheduled for use by 
that engine'' means all fuel in any tank intended for use by a specific 
engine.
    (ii) The fuel system design must clearly indicate the engine for 
which fuel in any tank is scheduled.
    (iii) Compliance with this paragraph must require no pilot action 
after completion of the engine starting phase of operations.
    (3) For single-engine airplanes, require no pilot action after 
completion of the engine starting phase of operations unless means are 
provided that unmistakenly alert the pilot to take any needed action at 
least five minutes prior to the needed action; such pilot action must 
not cause any change in engine operation; and such pilot action must not 
distract pilot attention from essential flight duties during any phase 
of operations for which the airplane is approved.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13093, Aug. 13, 1969; Amdt. 23-43, 58 FR 18971, Apr. 9, 1993; Amdt. 
23-51, 61 FR 5136, Feb. 9, 1996]



Sec. 23.957  Flow between interconnected tanks.

    (a) It must be impossible, in a gravity feed system with 
interconnected tank outlets, for enough fuel to flow between the tanks 
to cause an overflow of fuel from any tank vent under the

[[Page 276]]

conditions in Sec. 23.959, except that full tanks must be used.
    (b) If fuel can be pumped from one tank to another in flight, the 
fuel tank vents and the fuel transfer system must be designed so that no 
structural damage to any airplane component can occur because of 
overfilling of any tank.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18972, Apr. 9, 1993]



Sec. 23.959  Unusable fuel supply.

    (a) The unusable fuel supply for each tank must be established as 
not less than that quantity at which the first evidence of 
malfunctioning occurs under the most adverse fuel feed condition 
occurring under each intended operation and flight maneuver involving 
that tank. Fuel system component failures need not be considered.
    (b) The effect on the usable fuel quantity as a result of a failure 
of any pump shall be determined.

[Amdt. 23-7, 34 FR 13093, Aug. 13, 1969, as amended by Amdt. 23-18, 42 
FR 15041, Mar. 17, 1977; Amdt. 23-51, 61 FR 5136, Feb. 9, 1996]



Sec. 23.961  Fuel system hot weather operation.

    Each fuel system must be free from vapor lock when using fuel at its 
critical temperature, with respect to vapor formation, when operating 
the airplane in all critical operating and environmental conditions for 
which approval is requested. For turbine fuel, the initial temperature 
must be 110 [deg]F, -0[deg], +5 [deg]F or the maximum outside air 
temperature for which approval is requested, whichever is more critical.

[Doc. No. 26344, 58 FR 18972, Apr. 9, 1993; 58 FR 27060, May 6, 1993]



Sec. 23.963  Fuel tanks: General.

    (a) Each fuel tank must be able to withstand, without failure, the 
vibration, inertia, fluid, and structural loads that it may be subjected 
to in operation.
    (b) Each flexible fuel tank liner must be shown to be suitable for 
the particular application.
    (c) Each integral fuel tank must have adequate facilities for 
interior inspection and repair.
    (d) The total usable capacity of the fuel tanks must be enough for 
at least one-half hour of operation at maximum continuous power.
    (e) Each fuel quantity indicator must be adjusted, as specified in 
Sec. 23.1337(b), to account for the unusable fuel supply determined 
under Sec. 23.959(a).

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt 23-34, 52 FR 1832, Jan. 15, 1987; Amdt. 23-43, 58 FR 
18972, Apr. 9, 1993; Amdt. 23-51, 61 FR 5136, Feb. 9, 1996]



Sec. 23.965  Fuel tank tests.

    (a) Each fuel tank must be able to withstand the following pressures 
without failure or leakage:
    (1) For each conventional metal tank and nonmetallic tank with walls 
not supported by the airplane structure, a pressure of 3.5 p.s.i., or 
that pressure developed during maximum ultimate acceleration with a full 
tank, whichever is greater.
    (2) For each integral tank, the pressure developed during the 
maximum limit acceleration of the airplane with a full tank, with 
simultaneous application of the critical limit structural loads.
    (3) For each nonmetallic tank with walls supported by the airplane 
structure and constructed in an acceptable manner using acceptable basic 
tank material, and with actual or simulated support conditions, a 
pressure of 2 p.s.i. for the first tank of a specific design. The 
supporting structure must be designed for the critical loads occurring 
in the flight or landing strength conditions combined with the fuel 
pressure loads resulting from the corresponding accelerations.
    (b) Each fuel tank with large, unsupported, or unstiffened flat 
surfaces,whose failure or deformation could cause fuel leakage, must be 
able to withstand the following test without leakage, failure, or 
excessive deformation of the tank walls:
    (1) Each complete tank assembly and its support must be vibration 
tested while mounted to simulate the actual installation.
    (2) Except as specified in paragraph (b)(4) of this section, the 
tank assembly must be vibrated for 25 hours at a total displacement of 
not less than \1/32\ of an

[[Page 277]]

inch (unless another displacement is substantiated) while \2/3\ filled 
with water or other suitable test fluid.
    (3) The test frequency of vibration must be as follows:
    (i) If no frequency of vibration resulting from any rpm within the 
normal operating range of engine or propeller speeds is critical, the 
test frequency of vibration is:
    (A) The number of cycles per minute obtained by multiplying the 
maximum continuous propeller speed in rpm by 0.9 for propeller-driven 
airplanes, and
    (B) For non-propeller driven airplanes the test frequency of 
vibration is 2,000 cycles per minute.
    (ii) If only one frequency of vibration resulting from any rpm 
within the normal operating range of engine or propeller speeds is 
critical, that frequency of vibration must be the test frequency.
    (iii) If more than one frequency of vibration resulting from any rpm 
within the normal operating range of engine or propeller speeds is 
critical, the most critical of these frequencies must be the test 
frequency.
    (4) Under paragraph (b)(3) (ii) and (iii) of this section, the time 
of test must be adjusted to accomplish the same number of vibration 
cycles that would be accomplished in 25 hours at the frequency specified 
in paragraph (b)(3)(i) of this section.
    (5) During the test, the tank assembly must be rocked at a rate of 
16 to 20 complete cycles per minute, through an angle of 15[deg] on 
either side of the horizontal (30[deg] total), about an axis parallel to 
the axis of the fuselage, for 25 hours.
    (c) Each integral tank using methods of construction and sealing not 
previously proven to be adequate by test data or service experience must 
be able to withstand the vibration test specified in paragraphs (b)(1) 
through (4) of this section.
    (d) Each tank with a nonmetallic liner must be subjected to the 
sloshing test outlined in paragraph (b)(5) of this section, with the 
fuel at room temperature. In addition, a specimen liner of the same 
basic construction as that to be used in the airplane must, when 
installed in a suitable test tank, withstand the sloshing test with fuel 
at a temperature of 110 [deg]F.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18972, Apr. 9, 1993; Amdt. 23-43, 61 FR 253, Jan. 4, 1996; Amdt. 
23-51, 61 FR 5136, Feb. 9, 1996]



Sec. 23.967  Fuel tank installation.

    (a) Each fuel tank must be supported so that tank loads are not 
concentrated. In addition--
    (1) There must be pads, if necessary, to prevent chafing between 
each tank and its supports;
    (2) Padding must be nonabsorbent or treated to prevent the 
absorption of fuel;
    (3) If a flexible tank liner is used, it must be supported so that 
it is not required to withstand fluid loads;
    (4) Interior surfaces adjacent to the liner must be smooth and free 
from projections that could cause wear, unless--
    (i) Provisions are made for protection of the liner at those points; 
or
    (ii) The construction of the liner itself provides such protection; 
and
    (5) A positive pressure must be maintained within the vapor space of 
each bladder cell under any condition of operation, except for a 
particular condition for which it is shown that a zero or negative 
pressure will not cause the bladder cell to collapse; and
    (6) Syphoning of fuel (other than minor spillage) or collapse of 
bladder fuel cells may not result from improper securing or loss of the 
fuel filler cap.
    (b) Each tank compartment must be ventilated and drained to prevent 
the accumulation of flammable fluids or vapors. Each compartment 
adjacent to a tank that is an integral part of the airplane structure 
must also be ventilated and drained.
    (c) No fuel tank may be on the engine side of the firewall. There 
must be at least one-half inch of clearance between the fuel tank and 
the firewall. No part of the engine nacelle skin that lies immediately 
behind a major air opening from the engine compartment may act as the 
wall of an integral tank.
    (d) Each fuel tank must be isolated from personnel compartments by a 
fume-proof and fuel-proof enclosure

[[Page 278]]

that is vented and drained to the exterior of the airplane. The required 
enclosure must sustain any personnel compartment pressurization loads 
without permanent deformation or failure under the conditions of 
Sec. Sec. 23.365 and 23.843 of this part. A bladder-type fuel cell, if 
used, must have a retaining shell at least equivalent to a metal fuel 
tank in structural integrity.
    (e) Fuel tanks must be designed, located, and installed so as to 
retain fuel:
    (1) When subjected to the inertia loads resulting from the ultimate 
static load factors prescribed in Sec. 23.561(b)(2) of this part; and
    (2) Under conditions likely to occur when the airplane lands on a 
paved runway at a normal landing speed under each of the following 
conditions:
    (i) The airplane in a normal landing attitude and its landing gear 
retracted.
    (ii) The most critical landing gear leg collapsed and the other 
landing gear legs extended.

In showing compliance with paragraph (e)(2) of this section, the tearing 
away of an engine mount must be considered unless all the engines are 
installed above the wing or on the tail or fuselage of the airplane.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13903, Aug. 13, 1969; Amdt. 23-14, 38 FR 31823, Nov. 19, 1973; Amdt. 
23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-26, 45 FR 60171, Sept. 11, 
1980; Amdt. 23-36, 53 FR 30815, Aug. 15, 1988; Amdt. 23-43, 58 FR 18972, 
Apr. 9, 1993]



Sec. 23.969  Fuel tank expansion space.

    Each fuel tank must have an expansion space of not less than two 
percent of the tank capacity, unless the tank vent discharges clear of 
the airplane (in which case no expansion space is required). It must be 
impossible to fill the expansion space inadvertently with the airplane 
in the normal ground attitude.



Sec. 23.971  Fuel tank sump.

    (a) Each fuel tank must have a drainable sump with an effective 
capacity, in the normal ground and flight attitudes, of 0.25 percent of 
the tank capacity, or \1/16\ gallon, whichever is greater.
    (b) Each fuel tank must allow drainage of any hazardous quantity of 
water from any part of the tank to its sump with the airplane in the 
normal ground attitude.
    (c) Each reciprocating engine fuel system must have a sediment bowl 
or chamber that is accessible for drainage; has a capacity of 1 ounce 
for every 20 gallons of fuel tank capacity; and each fuel tank outlet is 
located so that, in the normal flight attitude, water will drain from 
all parts of the tank except the sump to the sediment bowl or chamber.
    (d) Each sump, sediment bowl, and sediment chamber drain required by 
paragraphs (a), (b), and (c) of this section must comply with the drain 
provisions of Sec. 23.999(b)(1) and (b)(2).

[Doc. No. 26344, 58 FR 18972, Apr. 9, 1993; 58 FR 27060, May 6, 1993]



Sec. 23.973  Fuel tank filler connection.

    (a) Each fuel tank filler connection must be marked as prescribed in 
Sec. 23.1557(c).
    (b) Spilled fuel must be prevented from entering the fuel tank 
compartment or any part of the airplane other than the tank itself.
    (c) Each filler cap must provide a fuel-tight seal for the main 
filler opening. However, there may be small openings in the fuel tank 
cap for venting purposes or for the purpose of allowing passage of a 
fuel gauge through the cap provided such openings comply with the 
requirements of Sec. 23.975(a).
    (d) Each fuel filling point, except pressure fueling connection 
points, must have a provision for electrically bonding the airplane to 
ground fueling equipment.
    (e) For airplanes with engines requiring gasoline as the only 
permissible fuel, the inside diameter of the fuel filler opening must be 
no larger than 2.36 inches.
    (f) For airplanes with turbine engines, the inside diameter of the 
fuel filler opening must be no smaller than 2.95 inches.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-43, 58 FR 
18972, Apr. 9, 1993; Amdt. 23-51, 61 FR 5136, Feb. 9, 1996]

[[Page 279]]



Sec. 23.975  Fuel tank vents and carburetor vapor vents.

    (a) Each fuel tank must be vented from the top part of the expansion 
space. In addition--
    (1) Each vent outlet must be located and constructed in a manner 
that minimizes the possibility of its being obstructed by ice or other 
foreign matter;
    (2) Each vent must be constructed to prevent siphoning of fuel 
during normal operation;
    (3) The venting capacity must allow the rapid relief of excessive 
differences of pressure between the interior and exterior of the tank;
    (4) Airspaces of tanks with interconnected outlets must be 
interconnected;
    (5) There may be no point in any vent line where moisture can 
accumulate with the airplane in either the ground or level flight 
attitudes, unless drainage is provided. Any drain valve installed must 
be accessible for drainage;
    (6) No vent may terminate at a point where the discharge of fuel 
from the vent outlet will constitute a fire hazard or from which fumes 
may enter personnel compartments; and
    (7) Vents must be arranged to prevent the loss of fuel, except fuel 
discharged because of thermal expansion, when the airplane is parked in 
any direction on a ramp having a one-percent slope.
    (b) Each carburetor with vapor elimination connections and each fuel 
injection engine employing vapor return provisions must have a separate 
vent line to lead vapors back to the top of one of the fuel tanks. If 
there is more than one tank and it is necessary to use these tanks in a 
definite sequence for any reason, the vapor vent line must lead back to 
the fuel tank to be used first, unless the relative capacities of the 
tanks are such that return to another tank is preferable.
    (c) For acrobatic category airplanes, excessive loss of fuel during 
acrobatic maneuvers, including short periods of inverted flight, must be 
prevented. It must be impossible for fuel to siphon from the vent when 
normal flight has been resumed after any acrobatic maneuver for which 
certification is requested.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-29, 49 FR 
6847, Feb. 23, 1984; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993; Amdt. 23-
51, 61 FR 5136, Feb. 9, 1996]



Sec. 23.977  Fuel tank outlet.

    (a) There must be a fuel strainer for the fuel tank outlet or for 
the booster pump. This strainer must--
    (1) For reciprocating engine powered airplanes, have 8 to 16 meshes 
per inch; and
    (2) For turbine engine powered airplanes, prevent the passage of any 
object that could restrict fuel flow or damage any fuel system 
component.
    (b) The clear area of each fuel tank outlet strainer must be at 
least five times the area of the outlet line.
    (c) The diameter of each strainer must be at least that of the fuel 
tank outlet.
    (d) Each strainer must be accessible for inspection and cleaning.

[Amdt. 23-17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23-43, 58 
FR 18973, Apr. 9, 1993]



Sec. 23.979  Pressure fueling systems.

    For pressure fueling systems, the following apply:
    (a) Each pressure fueling system fuel manifold connection must have 
means to prevent the escape of hazardous quantities of fuel from the 
system if the fuel entry valve fails.
    (b) An automatic shutoff means must be provided to prevent the 
quantity of fuel in each tank from exceeding the maximum quantity 
approved for that tank. This means must--
    (1) Allow checking for proper shutoff operation before each fueling 
of the tank; and
    (2) For commuter category airplanes, indicate at each fueling 
station, a failure of the shutoff means to stop the fuel flow at the 
maximum quantity approved for that tank.
    (c) A means must be provided to prevent damage to the fuel system in 
the event of failure of the automatic shutoff means prescribed in 
paragraph (b) of this section.

[[Page 280]]

    (d) All parts of the fuel system up to the tank which are subjected 
to fueling pressures must have a proof pressure of 1.33 times, and an 
ultimate pressure of at least 2.0 times, the surge pressure likely to 
occur during fueling.

[Amdt. 23-14, 38 FR 31823, Nov. 19, 1973, as amended by Amdt. 23-51, 61 
FR 5137, Feb. 9, 1996]

                         Fuel System Components



Sec. 23.991  Fuel pumps.

    (a) Main pumps. For main pumps, the following apply:
    (1) For reciprocating engine installations having fuel pumps to 
supply fuel to the engine, at least one pump for each engine must be 
directly driven by the engine and must meet Sec. 23.955. This pump is a 
main pump.
    (2) For turbine engine installations, each fuel pump required for 
proper engine operation, or required to meet the fuel system 
requirements of this subpart (other than those in paragraph (b) of this 
section), is a main pump. In addition--
    (i) There must be at least one main pump for each turbine engine;
    (ii) The power supply for the main pump for each engine must be 
independent of the power supply for each main pump for any other engine; 
and
    (iii) For each main pump, provision must be made to allow the bypass 
of each positive displacement fuel pump other than a fuel injection pump 
approved as part of the engine.
    (b) Emergency pumps. There must be an emergency pump immediately 
available to supply fuel to the engine if any main pump (other than a 
fuel injection pump approved as part of an engine) fails. The power 
supply for each emergency pump must be independent of the power supply 
for each corresponding main pump.
    (c) Warning means. If both the main pump and emergency pump operate 
continuously, there must be a means to indicate to the appropriate 
flight crewmembers a malfunction of either pump.
    (d) Operation of any fuel pump may not affect engine operation so as 
to create a hazard, regardless of the engine power or thrust setting or 
the functional status of any other fuel pump.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13093, Aug. 13, 1969; Amdt. 23-26, 45 FR 60171, Sept. 11, 1980; Amdt. 
23-43, 58 FR 18973, Apr. 9, 1993]



Sec. 23.993  Fuel system lines and fittings.

    (a) Each fuel line must be installed and supported to prevent 
excessive vibration and to withstand loads due to fuel pressure and 
accelerated flight conditions.
    (b) Each fuel line connected to components of the airplane between 
which relative motion could exist must have provisions for flexibility.
    (c) Each flexible connection in fuel lines that may be under 
pressure and subjected to axial loading must use flexible hose 
assemblies.
    (d) Each flexible hose must be shown to be suitable for the 
particular application.
    (e) No flexible hose that might be adversely affected by exposure to 
high temperatures may be used where excessive temperatures will exist 
during operation or after engine shutdown.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18973, Apr. 9, 1993]



Sec. 23.994  Fuel system components.

    Fuel system components in an engine nacelle or in the fuselage must 
be protected from damage which could result in spillage of enough fuel 
to constitute a fire hazard as a result of a wheels-up landing on a 
paved runway.

[Amdt. 23-29, 49 FR 6847, Feb. 23, 1984]



Sec. 23.995  Fuel valves and controls.

    (a) There must be a means to allow appropriate flight crew members 
to rapidly shut off, in flight, the fuel to each engine individually.
    (b) No shutoff valve may be on the engine side of any firewall. In 
addition, there must be means to--
    (1) Guard against inadvertent operation of each shutoff valve; and
    (2) Allow appropriate flight crew members to reopen each valve 
rapidly after it has been closed.

[[Page 281]]

    (c) Each valve and fuel system control must be supported so that 
loads resulting from its operation or from accelerated flight conditions 
are not transmitted to the lines connected to the valve.
    (d) Each valve and fuel system control must be installed so that 
gravity and vibration will not affect the selected position.
    (e) Each fuel valve handle and its connections to the valve 
mechanism must have design features that minimize the possibility of 
incorrect installation.
    (f) Each check valve must be constructed, or otherwise incorporate 
provisions, to preclude incorrect assembly or connection of the valve.
    (g) Fuel tank selector valves must--
    (1) Require a separate and distinct action to place the selector in 
the ``OFF'' position; and
    (2) Have the tank selector positions located in such a manner that 
it is impossible for the selector to pass through the ``OFF'' position 
when changing from one tank to another.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31823, Nov. 19, 1973; Amdt. 23-17, 41 FR 55465, Dec. 20, 1976; 
Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-29, 49 FR 6847, Feb. 
23, 1984]



Sec. 23.997  Fuel strainer or filter.

    There must be a fuel strainer or filter between the fuel tank outlet 
and the inlet of either the fuel metering device or an engine driven 
positive displacement pump, whichever is nearer the fuel tank outlet. 
This fuel strainer or filter must--
    (a) Be accessible for draining and cleaning and must incorporate a 
screen or element which is easily removable;
    (b) Have a sediment trap and drain except that it need not have a 
drain if the strainer or filter is easily removable for drain purposes;
    (c) Be mounted so that its weight is not supported by the connecting 
lines or by the inlet or outlet connections of the strainer or filter 
itself, unless adequate strength margins under all loading conditions 
are provided in the lines and connections; and
    (d) Have the capacity (with respect to operating limitations 
established for the engine) to ensure that engine fuel system 
functioning is not impaired, with the fuel contaminated to a degree 
(with respect to particle size and density) that is greater than that 
established for the engine during its type certification.
    (e) In addition, for commuter category airplanes, unless means are 
provided in the fuel system to prevent the accumulation of ice on the 
filter, a means must be provided to automatically maintain the fuel flow 
if ice clogging of the filter occurs.

[Amdt. 23-15, 39 FR 35459, Oct. 1, 1974, as amended by Amdt. 23-29, 49 
FR 6847, Feb. 23, 1984; Amdt. 23-34, 52 FR 1832, Jan. 15, 1987; Amdt. 
23-43, 58 FR 18973, Apr. 9, 1993]



Sec. 23.999  Fuel system drains.

    (a) There must be at least one drain to allow safe drainage of the 
entire fuel system with the airplane in its normal ground attitude.
    (b) Each drain required by paragraph (a) of this section and Sec. 
23.971 must--
    (1) Discharge clear of all parts of the airplane;
    (2) Have a drain valve--
    (i) That has manual or automatic means for positive locking in the 
closed position;
    (ii) That is readily accessible;
    (iii) That can be easily opened and closed;
    (iv) That allows the fuel to be caught for examination;
    (v) That can be observed for proper closing; and
    (vi) That is either located or protected to prevent fuel spillage in 
the event of a landing with landing gear retracted.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 
41 FR 55465, Dec. 20, 1976; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]



Sec. 23.1001  Fuel jettisoning system.

    (a) If the design landing weight is less than that permitted under 
the requirements of Sec. 23.473(b), the airplane must have a fuel 
jettisoning system installed that is able to jettison enough fuel to 
bring the maximum weight down to the design landing weight. The average 
rate of fuel jettisoning must be at least 1 percent of the maximum 
weight per minute, except that the time required to jettison the fuel 
need not be less than 10 minutes.

[[Page 282]]

    (b) Fuel jettisoning must be demonstrated at maximum weight with 
flaps and landing gear up and in--
    (1) A power-off glide at 1.4 VS1;
    (2) A climb, at the speed at which the one-engine-inoperative 
enroute climb data have been established in accordance with Sec. 
23.69(b), with the critical engine inoperative and the remaining engines 
at maximum continuous power; and
    (3) Level flight at 1.4 VS1, if the results of the tests 
in the conditions specified in paragraphs (b)(1) and (2) of this section 
show that this condition could be critical.
    (c) During the flight tests prescribed in paragraph (b) of this 
section, it must be shown that--
    (1) The fuel jettisoning system and its operation are free from fire 
hazard;
    (2) The fuel discharges clear of any part of the airplane;
    (3) Fuel or fumes do not enter any parts of the airplane; and
    (4) The jettisoning operation does not adversely affect the 
controllability of the airplane.
    (d) For reciprocating engine powered airplanes, the jettisoning 
system must be designed so that it is not possible to jettison the fuel 
in the tanks used for takeoff and landing below the level allowing 45 
minutes flight at 75 percent maximum continuous power. However, if there 
is an auxiliary control independent of the main jettisoning control, the 
system may be designed to jettison all the fuel.
    (e) For turbine engine powered airplanes, the jettisoning system 
must be designed so that it is not possible to jettison fuel in the 
tanks used for takeoff and landing below the level allowing climb from 
sea level to 10,000 feet and thereafter allowing 45 minutes cruise at a 
speed for maximum range.
    (f) The fuel jettisoning valve must be designed to allow flight 
crewmembers to close the valve during any part of the jettisoning 
operation.
    (g) Unless it is shown that using any means (including flaps, slots, 
and slats) for changing the airflow across or around the wings does not 
adversely affect fuel jettisoning, there must be a placard, adjacent to 
the jettisoning control, to warn flight crewmembers against jettisoning 
fuel while the means that change the airflow are being used.
    (h) The fuel jettisoning system must be designed so that any 
reasonably probable single malfunction in the system will not result in 
a hazardous condition due to unsymmetrical jettisoning of, or inability 
to jettison, fuel.

[Amdt. 23-7, 34 FR 13094, Aug. 13, 1969, as amended by Amdt. 23-43, 58 
FR 18973, Apr. 9, 1993; Amdt. 23-51, 61 FR 5137, Feb. 9, 1996]

                               Oil System



Sec. 23.1011  General.

    (a) For oil systems and components that have been approved under the 
engine airworthiness requirements and where those requirements are equal 
to or more severe than the corresponding requirements of subpart E of 
this part, that approval need not be duplicated. Where the requirements 
of subpart E of this part are more severe, substantiation must be shown 
to the requirements of subpart E of this part.
    (b) Each engine must have an independent oil system that can supply 
it with an appropriate quantity of oil at a temperature not above that 
safe for continuous operation.
    (c) The usable oil tank capacity may not be less than the product of 
the endurance of the airplane under critical operating conditions and 
the maximum oil consumption of the engine under the same conditions, 
plus a suitable margin to ensure adequate circulation and cooling.
    (d) For an oil system without an oil transfer system, only the 
usable oil tank capacity may be considered. The amount of oil in the 
engine oil lines, the oil radiator, and the feathering reserve, may not 
be considered.
    (e) If an oil transfer system is used, and the transfer pump can 
pump some of the oil in the transfer lines into the main engine oil 
tanks, the amount of oil in these lines that can be pumped by the 
transfer pump may be included in the oil capacity.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18973, Apr. 9, 1993]



Sec. 23.1013  Oil tanks.

    (a) Installation. Each oil tank must be installed to--

[[Page 283]]

    (1) Meet the requirements of Sec. 23.967 (a) and (b); and
    (2) Withstand any vibration, inertia, and fluid loads expected in 
operation.
    (b) Expansion space. Oil tank expansion space must be provided so 
that--
    (1) Each oil tank used with a reciprocating engine has an expansion 
space of not less than the greater of 10 percent of the tank capacity or 
0.5 gallon, and each oil tank used with a turbine engine has an 
expansion space of not less than 10 percent of the tank capacity; and
    (2) It is impossible to fill the expansion space inadvertently with 
the airplane in the normal ground attitude.
    (c) Filler connection. Each oil tank filler connection must be 
marked as specified in Sec. 23.1557(c). Each recessed oil tank filler 
connection of an oil tank used with a turbine engine, that can retain 
any appreciable quantity of oil, must have provisions for fitting a 
drain.
    (d) Vent. Oil tanks must be vented as follows:
    (1) Each oil tank must be vented to the engine from the top part of 
the expansion space so that the vent connection is not covered by oil 
under any normal flight condition.
    (2) Oil tank vents must be arranged so that condensed water vapor 
that might freeze and obstruct the line cannot accumulate at any point.
    (3) For acrobatic category airplanes, there must be means to prevent 
hazardous loss of oil during acrobatic maneuvers, including short 
periods of inverted flight.
    (e) Outlet. No oil tank outlet may be enclosed by any screen or 
guard that would reduce the flow of oil below a safe value at any 
operating temperature. No oil tank outlet diameter may be less than the 
diameter of the engine oil pump inlet. Each oil tank used with a turbine 
engine must have means to prevent entrance into the tank itself, or into 
the tank outlet, of any object that might obstruct the flow of oil 
through the system. There must be a shutoff valve at the outlet of each 
oil tank used with a turbine engine, unless the external portion of the 
oil system (including oil tank supports) is fireproof.
    (f) Flexible liners. Each flexible oil tank liner must be of an 
acceptable kind.
    (g) Each oil tank filler cap of an oil tank that is used with an 
engine must provide an oiltight seal.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-15, 
39 FR 35459 Oct. 1, 1974; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993; Amdt. 
23-51, 61 FR 5137, Feb. 9, 1996]



Sec. 23.1015  Oil tank tests.

    Each oil tank must be tested under Sec. 23.965, except that--
    (a) The applied pressure must be five p.s.i. for the tank 
construction instead of the pressures specified in Sec. 23.965(a);
    (b) For a tank with a nonmetallic liner the test fluid must be oil 
rather than fuel as specified in Sec. 23.965(d), and the slosh test on 
a specimen liner must be conducted with the oil at 250 [deg]F.; and
    (c) For pressurized tanks used with a turbine engine, the test 
pressure may not be less than 5 p.s.i. plus the maximum operating 
pressure of the tank.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-15, 
39 FR 35460, Oct. 1, 1974]



Sec. 23.1017  Oil lines and fittings.

    (a) Oil lines. Oil lines must meet Sec. 23.993 and must accommodate 
a flow of oil at a rate and pressure adequate for proper engine 
functioning under any normal operating condition.
    (b) Breather lines. Breather lines must be arranged so that--
    (1) Condensed water vapor or oil that might freeze and obstruct the 
line cannot accumulate at any point;
    (2) The breather discharge will not constitute a fire hazard if 
foaming occurs, or cause emitted oil to strike the pilot's windshield;
    (3) The breather does not discharge into the engine air induction 
system; and
    (4) For acrobatic category airplanes, there is no excessive loss of 
oil from the breather during acrobatic maneuvers, including short 
periods of inverted flight.

[[Page 284]]

    (5) The breather outlet is protected against blockage by ice or 
foreign matter.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13094, Aug. 13, 1969; Amdt. 23-14, 38 FR 31823, Nov. 19, 1973]



Sec. 23.1019  Oil strainer or filter.

    (a) Each turbine engine installation must incorporate an oil 
strainer or filter through which all of the engine oil flows and which 
meets the following requirements:
    (1) Each oil strainer or filter that has a bypass, must be 
constructed and installed so that oil will flow at the normal rate 
through the rest of the system with the strainer or filter completely 
blocked.
    (2) The oil strainer or filter must have the capacity (with respect 
to operating limitations established for the engine) to ensure that 
engine oil system functioning is not impaired when the oil is 
contaminated to a degree (with respect to particle size and density) 
that is greater than that established for the engine for its type 
certification.
    (3) The oil strainer or filter, unless it is installed at an oil 
tank outlet, must incorporate a means to indicate contamination before 
it reaches the capacity established in accordance with paragraph (a)(2) 
of this section.
    (4) The bypass of a strainer or filter must be constructed and 
installed so that the release of collected contaminants is minimized by 
appropriate location of the bypass to ensure that collected contaminants 
are not in the bypass flow path.
    (5) An oil strainer or filter that has no bypass, except one that is 
installed at an oil tank outlet, must have a means to connect it to the 
warning system required in Sec. 23.1305(c)(9).
    (b) Each oil strainer or filter in a powerplant installation using 
reciprocating engines must be constructed and installed so that oil will 
flow at the normal rate through the rest of the system with the strainer 
or filter element completely blocked.

[Amdt. 23-15, 39 FR 35460, Oct. 1, 1974, as amended by Amdt. 23-29, 49 
FR 6847, Feb. 23, 1984; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]



Sec. 23.1021  Oil system drains.

    A drain (or drains) must be provided to allow safe drainage of the 
oil system. Each drain must--
    (a) Be accessible;
    (b) Have drain valves, or other closures, employing manual or 
automatic shut-off means for positive locking in the closed position; 
and
    (c) Be located or protected to prevent inadvertent operation.

[Amdt. 23-29, 49 FR 6847, Feb. 23, 1984, as amended by Amdt. 23-43, 58 
FR 18973, Apr. 9, 1993]



Sec. 23.1023  Oil radiators.

    Each oil radiator and its supporting structures must be able to 
withstand the vibration, inertia, and oil pressure loads to which it 
would be subjected in operation.



Sec. 23.1027  Propeller feathering system.

    (a) If the propeller feathering system uses engine oil and that oil 
supply can become depleted due to failure of any part of the oil system, 
a means must be incorporated to reserve enough oil to operate the 
feathering system.
    (b) The amount of reserved oil must be enough to accomplish 
feathering and must be available only to the feathering pump.
    (c) The ability of the system to accomplish feathering with the 
reserved oil must be shown.
    (d) Provision must be made to prevent sludge or other foreign matter 
from affecting the safe operation of the propeller feathering system.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31823, Nov. 19, 1973; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]

                                 Cooling



Sec. 23.1041  General.

    The powerplant and auxiliary power unit cooling provisions must 
maintain the temperatures of powerplant components and engine fluids, 
and auxiliary power unit components and fluids within the limits 
established for those components and fluids under the most adverse 
ground, water, and flight operations to the maximum altitude and maximum 
ambient atmospheric temperature conditions for which approval

[[Page 285]]

is requested, and after normal engine and auxiliary power unit shutdown.

[Doc. No. 26344, 58 FR 18973, Apr. 9, 1993, as amended by Amdt. 23-51, 
61 FR 5137, Feb. 9, 1996]



Sec. 23.1043  Cooling tests.

    (a) General. Compliance with Sec. 23.1041 must be shown on the 
basis of tests, for which the following apply:
    (1) If the tests are conducted under ambient atmospheric temperature 
conditions deviating from the maximum for which approval is requested, 
the recorded powerplant temperatures must be corrected under paragraphs 
(c) and (d) of this section, unless a more rational correction method is 
applicable.
    (2) No corrected temperature determined under paragraph (a)(1) of 
this section may exceed established limits.
    (3) The fuel used during the cooling tests must be of the minimum 
grade approved for the engine.
    (4) For turbocharged engines, each turbocharger must be operated 
through that part of the climb profile for which operation with the 
turbocharger is requested.
    (5) For a reciprocating engine, the mixture settings must be the 
leanest recommended for climb.
    (b) Maximum ambient atmospheric temperature. A maximum ambient 
atmospheric temperature corresponding to sea level conditions of at 
least 100 degrees F must be established. The assumed temperature lapse 
rate is 3.6 degrees F per thousand feet of altitude above sea level 
until a temperature of -69.7 degrees F is reached, above which altitude 
the temperature is considered constant at -69.7 degrees F. However, for 
winterization installations, the applicant may select a maximum ambient 
atmospheric temperature corresponding to sea level conditions of less 
than 100 degrees F.
    (c) Correction factor (except cylinder barrels). Temperatures of 
engine fluids and powerplant components (except cylinder barrels) for 
which temperature limits are established, must be corrected by adding to 
them the difference between the maximum ambient atmospheric temperature 
for the relevant altitude for which approval has been requested and the 
temperature of the ambient air at the time of the first occurrence of 
the maximum fluid or component temperature recorded during the cooling 
test.
    (d) Correction factor for cylinder barrel temperatures. Cylinder 
barrel temperatures must be corrected by adding to them 0.7 times the 
difference between the maximum ambient atmospheric temperature for the 
relevant altitude for which approval has been requested and the 
temperature of the ambient air at the time of the first occurrence of 
the maximum cylinder barrel temperature recorded during the cooling 
test.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13094, Aug. 13, 1969; Amdt. 23-21, 43 FR 2319, Jan. 16, 1978; Amdt. 
23-51, 61 FR 5137, Feb. 9, 1996]



Sec. 23.1045  Cooling test procedures for turbine engine powered
airplanes.

    (a) Compliance with Sec. 23.1041 must be shown for all phases of 
operation. The airplane must be flown in the configurations, at the 
speeds, and following the procedures recommended in the Airplane Flight 
Manual for the relevant stage of flight, that correspond to the 
applicable performance requirements that are critical to cooling.
    (b) Temperatures must be stabilized under the conditions from which 
entry is made into each stage of flight being investigated, unless the 
entry condition normally is not one during which component and engine 
fluid temperatures would stabilize (in which case, operation through the 
full entry condition must be conducted before entry into the stage of 
flight being investigated in order to allow temperatures to reach their 
natural levels at the time of entry). The takeoff cooling test must be 
preceded by a period during which the powerplant component and engine 
fluid temperatures are stabilized with the engines at ground idle.
    (c) Cooling tests for each stage of flight must be continued until--
    (1) The component and engine fluid temperatures stabilize;
    (2) The stage of flight is completed; or
    (3) An operating limitation is reached.

[Amdt. 23-7, 34 FR 13094, Aug. 13, 1969, as amended by Amdt. 23-51, 61 
FR 5137, Feb. 9, 1996]

[[Page 286]]



Sec. 23.1047  Cooling test procedures for reciprocating engine powered
airplanes.

    Compliance with Sec. 23.1041 must be shown for the climb (or, for 
multiengine airplanes with negative one-engine-inoperative rates of 
climb, the descent) stage of flight. The airplane must be flown in the 
configurations, at the speeds and following the procedures recommended 
in the Airplane Flight Manual, that correspond to the applicable 
performance requirements that are critical to cooling.

[Amdt. 23-51, 61 FR 5137, Feb. 9, 1996]

                             Liquid Cooling



Sec. 23.1061  Installation.

    (a) General. Each liquid-cooled engine must have an independent 
cooling system (including coolant tank) installed so that--
    (1) Each coolant tank is supported so that tank loads are 
distributed over a large part of the tank surface;
    (2) There are pads or other isolation means between the tank and its 
supports to prevent chafing.
    (3) Pads or any other isolation means that is used must be 
nonabsorbent or must be treated to prevent absorption of flammable 
fluids; and
    (4) No air or vapor can be trapped in any part of the system, except 
the coolant tank expansion space, during filling or during operation.
    (b) Coolant tank. The tank capacity must be at least one gallon, 
plus 10 percent of the cooling system capacity. In addition--
    (1) Each coolant tank must be able to withstand the vibration, 
inertia, and fluid loads to which it may be subjected in operation;
    (2) Each coolant tank must have an expansion space of at least 10 
percent of the total cooling system capacity; and
    (3) It must be impossible to fill the expansion space inadvertently 
with the airplane in the normal ground attitude.
    (c) Filler connection. Each coolant tank filler connection must be 
marked as specified in Sec. 23.1557(c). In addition--
    (1) Spilled coolant must be prevented from entering the coolant tank 
compartment or any part of the airplane other than the tank itself; and
    (2) Each recessed coolant filler connection must have a drain that 
discharges clear of the entire airplane.
    (d) Lines and fittings. Each coolant system line and fitting must 
meet the requirements of Sec. 23.993, except that the inside diameter 
of the engine coolant inlet and outlet lines may not be less than the 
diameter of the corresponding engine inlet and outlet connections.
    (e) Radiators. Each coolant radiator must be able to withstand any 
vibration, inertia, and coolant pressure load to which it may normally 
be subjected. In addition--
    (1) Each radiator must be supported to allow expansion due to 
operating temperatures and prevent the transmittal of harmful vibration 
to the radiator; and
    (2) If flammable coolant is used, the air intake duct to the coolant 
radiator must be located so that (in case of fire) flames from the 
nacelle cannot strike the radiator.
    (f) Drains. There must be an accessible drain that--
    (1) Drains the entire cooling system (including the coolant tank, 
radiator, and the engine) when the airplane is in the normal ground 
altitude;
    (2) Discharges clear of the entire airplane; and
    (3) Has means to positively lock it closed.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18973, Apr. 9, 1993]



Sec. 23.1063  Coolant tank tests.

    Each coolant tank must be tested under Sec. 23.965, except that--
    (a) The test required by Sec. 23.965(a)(1) must be replaced with a 
similar test using the sum of the pressure developed during the maximum 
ultimate acceleration with a full tank or a pressure of 3.5 pounds per 
square inch, whichever is greater, plus the maximum working pressure of 
the system; and
    (b) For a tank with a nonmetallic liner the test fluid must be 
coolant rather than fuel as specified in Sec. 23.965(d), and the slosh 
test on a specimen liner must be conducted with the coolant at operating 
temperature.

[[Page 287]]

                            Induction System



Sec. 23.1091  Air induction system.

    (a) The air induction system for each engine and auxiliary power 
unit and their accessories must supply the air required by that engine 
and auxiliary power unit and their accessories under the operating 
conditions for which certification is requested.
    (b) Each reciprocating engine installation must have at least two 
separate air intake sources and must meet the following:
    (1) Primary air intakes may open within the cowling if that part of 
the cowling is isolated from the engine accessory section by a fire-
resistant diaphragm or if there are means to prevent the emergence of 
backfire flames.
    (2) Each alternate air intake must be located in a sheltered 
position and may not open within the cowling if the emergence of 
backfire flames will result in a hazard.
    (3) The supplying of air to the engine through the alternate air 
intake system may not result in a loss of excessive power in addition to 
the power loss due to the rise in air temperature.
    (4) Each automatic alternate air door must have an override means 
accessible to the flight crew.
    (5) Each automatic alternate air door must have a means to indicate 
to the flight crew when it is not closed.
    (c) For turbine engine powered airplanes--
    (1) There must be means to prevent hazardous quantities of fuel 
leakage or overflow from drains, vents, or other components of flammable 
fluid systems from entering the engine intake system; and
    (2) The airplane must be designed to prevent water or slush on the 
runway, taxiway, or other airport operating surfaces from being directed 
into the engine or auxiliary power unit air intake ducts in hazardous 
quantities. The air intake ducts must be located or protected so as to 
minimize the hazard of ingestion of foreign matter during takeoff, 
landing, and taxiing.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13095, Aug. 13, 1969; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993; 58 FR 
27060, May 6, 1993; Amdt. 23-51, 61 FR 5137, Feb. 9, 1996]



Sec. 23.1093  Induction system icing protection.

    (a) Reciprocating engines. Each reciprocating engine air induction 
system must have means to prevent and eliminate icing. Unless this is 
done by other means, it must be shown that, in air free of visible 
moisture at a temperature of 30 [deg]F--
    (1) Each airplane with sea level engines using conventional venturi 
carburetors has a preheater that can provide a heat rise of 90 [deg]F. 
with the engines at 75 percent of maximum continuous power;
    (2) Each airplane with altitude engines using conventional venturi 
carburetors has a preheater that can provide a heat rise of 120 [deg]F. 
with the engines at 75 percent of maximum continuous power;
    (3) Each airplane with altitude engines using fuel metering device 
tending to prevent icing has a preheater that, with the engines at 60 
percent of maximum continuous power, can provide a heat rise of--
    (i) 100 [deg]F.; or
    (ii) 40 [deg]F., if a fluid deicing system meeting the requirements 
of Sec. Sec. 23.1095 through 23.1099 is installed;
    (4) Each airplane with sea level engine(s) using fuel metering 
device tending to prevent icing has a sheltered alternate source of air 
with a preheat of not less than 60 [deg]F with the engines at 75 percent 
of maximum continuous power;
    (5) Each airplane with sea level or altitude engine(s) using fuel 
injection systems having metering components on which impact ice may 
accumulate has a preheater capable of providing a heat rise of 75 [deg]F 
when the engine is operating at 75 percent of its maximum continuous 
power; and
    (6) Each airplane with sea level or altitude engine(s) using fuel 
injection systems not having fuel metering components projecting into 
the airstream on which ice may form, and introducing fuel into the air 
induction system downstream of any components or other obstruction on 
which ice produced by fuel evaporation may form, has a sheltered 
alternate source of air with a preheat of not less than 60 [deg]F with 
the engines at 75 percent of its maximum continuous power.

[[Page 288]]

    (b) Turbine engines. (1) Each turbine engine and its air inlet 
system must operate throughout the flight power range of the engine 
(including idling), without the accumulation of ice on engine or inlet 
system components that would adversely affect engine operation or cause 
a serious loss of power or thrust--
    (i) Under the icing conditions specified in appendix C of part 25 of 
this chapter; and
    (ii) In snow, both falling and blowing, within the limitations 
established for the airplane for such operation.
    (2) Each turbine engine must idle for 30 minutes on the ground, with 
the air bleed available for engine icing protection at its critical 
condition, without adverse effect, in an atmosphere that is at a 
temperature between 15[deg] and 30 [deg]F (between -9[deg] and -1 
[deg]C) and has a liquid water content not less than 0.3 grams per cubic 
meter in the form of drops having a mean effective diameter not less 
than 20 microns, followed by momentary operation at takeoff power or 
thrust. During the 30 minutes of idle operation, the engine may be run 
up periodically to a moderate power or thrust setting in a manner 
acceptable to the Administrator.
    (c) Reciprocating engines with Superchargers. For airplanes with 
reciprocating engines having superchargers to pressurize the air before 
it enters the fuel metering device, the heat rise in the air caused by 
that supercharging at any altitude may be utilized in determining 
compliance with paragraph (a) of this section if the heat rise utilized 
is that which will be available, automatically, for the applicable 
altitudes and operating condition because of supercharging.

[Amdt. 23-7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23-15, 39 
FR 35460, Oct. 1, 1974; Amdt. 23-17, 41 FR 55465, Dec. 20, 1976; Amdt. 
23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-29, 49 FR 6847, Feb. 23, 
1984; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993; Amdt. 23-51, 61 FR 5137, 
Feb. 9, 1996]



Sec. 23.1095  Carburetor deicing fluid flow rate.

    (a) If a carburetor deicing fluid system is used, it must be able to 
simultaneously supply each engine with a rate of fluid flow, expressed 
in pounds per hour, of not less than 2.5 times the square root of the 
maximum continuous power of the engine.
    (b) The fluid must be introduced into the air induction system--
    (1) Close to, and upstream of, the carburetor; and
    (2) So that it is equally distributed over the entire cross section 
of the induction system air passages.



Sec. 23.1097  Carburetor deicing fluid system capacity.

    (a) The capacity of each carburetor deicing fluid system--
    (1) May not be less than the greater of--
    (i) That required to provide fluid at the rate specified in Sec. 
23.1095 for a time equal to three percent of the maximum endurance of 
the airplane; or
    (ii) 20 minutes at that flow rate; and
    (2) Need not exceed that required for two hours of operation.
    (b) If the available preheat exceeds 50 [deg]F. but is less than 100 
[deg]F., the capacity of the system may be decreased in proportion to 
the heat rise available in excess of 50 [deg]F.



Sec. 23.1099  Carburetor deicing fluid system detail design.

    Each carburetor deicing fluid system must meet the applicable 
requirements for the design of a fuel system, except as specified in 
Sec. Sec. 23.1095 and 23.1097.



Sec. 23.1101  Induction air preheater design.

    Each exhaust-heated, induction air preheater must be designed and 
constructed to--
    (a) Ensure ventilation of the preheater when the induction air 
preheater is not being used during engine operation;
    (b) Allow inspection of the exhaust manifold parts that it 
surrounds; and
    (c) Allow inspection of critical parts of the preheater itself.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18974, Apr. 9, 1993]



Sec. 23.1103  Induction system ducts.

    (a) Each induction system duct must have a drain to prevent the 
accumulation of fuel or moisture in the normal ground and flight 
attitudes. No drain

[[Page 289]]

may discharge where it will cause a fire hazard.
    (b) Each duct connected to components between which relative motion 
could exist must have means for flexibility.
    (c) Each flexible induction system duct must be capable of 
withstanding the effects of temperature extremes, fuel, oil, water, and 
solvents to which it is expected to be exposed in service and 
maintenance without hazardous deterioration or delamination.
    (d) For reciprocating engine installations, each induction system 
duct must be--
    (1) Strong enough to prevent induction system failures resulting 
from normal backfire conditions; and
    (2) Fire resistant in any compartment for which a fire extinguishing 
system is required.
    (e) Each inlet system duct for an auxiliary power unit must be--
    (1) Fireproof within the auxiliary power unit compartment;
    (2) Fireproof for a sufficient distance upstream of the auxiliary 
power unit compartment to prevent hot gas reverse flow from burning 
through the duct and entering any other compartment of the airplane in 
which a hazard would be created by the entry of the hot gases;
    (3) Constructed of materials suitable to the environmental 
conditions expected in service, except in those areas requiring 
fireproof or fire resistant materials; and
    (4) Constructed of materials that will not absorb or trap hazardous 
quantities of flammable fluids that could be ignited by a surge or 
reverse-flow condition.
    (f) Induction system ducts that supply air to a cabin pressurization 
system must be suitably constructed of material that will not produce 
hazardous quantities of toxic gases or isolated to prevent hazardous 
quantities of toxic gases from entering the cabin during a powerplant 
fire.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13095, Aug. 13, 1969; Amdt. 23-43, 58 FR 18974, Apr. 9, 1993]



Sec. 23.1105  Induction system screens.

    If induction system screens are used--
    (a) Each screen must be upstream of the carburetor or fuel injection 
system.
    (b) No screen may be in any part of the induction system that is the 
only passage through which air can reach the engine, unless--
    (1) The available preheat is at least 100 [deg]F.; and
    (2) The screen can be deiced by heated air;
    (c) No screen may be deiced by alcohol alone; and
    (d) It must be impossible for fuel to strike any screen.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1996, as 
amended by Amdt. 23-51, 61 FR 5137, Feb. 9, 1996]



Sec. 23.1107  Induction system filters.

    If an air filter is used to protect the engine against foreign 
material particles in the induction air supply--
    (a) Each air filter must be capable of withstanding the effects of 
temperature extremes, rain, fuel, oil, and solvents to which it is 
expected to be exposed in service and maintenance; and
    (b) Each air filter shall have a design feature to prevent material 
separated from the filter media from interfering with proper fuel 
metering operation.

[Doc. No. 26344, 58 FR 18974, Apr. 9, 1993, as amended by Amdt. 23-51, 
61 FR 5137, Feb. 9, 1996]



Sec. 23.1109  Turbocharger bleed air system.

    The following applies to turbocharged bleed air systems used for 
cabin pressurization:
    (a) The cabin air system may not be subject to hazardous 
contamination following any probable failure of the turbocharger or its 
lubrication system.
    (b) The turbocharger supply air must be taken from a source where it 
cannot be contaminated by harmful or hazardous gases or vapors following 
any probable failure or malfunction of the engine exhaust, hydraulic, 
fuel, or oil system.

[Amdt. 23-42, 56 FR 354, Jan. 3, 1991]



Sec. 23.1111  Turbine engine bleed air system.

    For turbine engine bleed air systems, the following apply:

[[Page 290]]

    (a) No hazard may result if duct rupture or failure occurs anywhere 
between the engine port and the airplane unit served by the bleed air.
    (b) The effect on airplane and engine performance of using maximum 
bleed air must be established.
    (c) Hazardous contamination of cabin air systems may not result from 
failures of the engine lubricating system.

[Amdt. 23-7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23-17, 41 
FR 55465, Dec. 20, 1976]

                             Exhaust System



Sec. 23.1121  General.

    For powerplant and auxiliary power unit installations, the following 
apply--
    (a) Each exhaust system must ensure safe disposal of exhaust gases 
without fire hazard or carbon monoxide contamination in any personnel 
compartment.
    (b) Each exhaust system part with a surface hot enough to ignite 
flammable fluids or vapors must be located or shielded so that leakage 
from any system carrying flammable fluids or vapors will not result in a 
fire caused by impingement of the fluids or vapors on any part of the 
exhaust system including shields for the exhaust system.
    (c) Each exhaust system must be separated by fireproof shields from 
adjacent flammable parts of the airplane that are outside of the engine 
and auxiliary power unit compartments.
    (d) No exhaust gases may discharge dangerously near any fuel or oil 
system drain.
    (e) No exhaust gases may be discharged where they will cause a glare 
seriously affecting pilot vision at night.
    (f) Each exhaust system component must be ventilated to prevent 
points of excessively high temperature.
    (g) If significant traps exist, each turbine engine and auxiliary 
power unit exhaust system must have drains discharging clear of the 
airplane, in any normal ground and flight attitude, to prevent fuel 
accumulation after the failure of an attempted engine or auxiliary power 
unit start.
    (h) Each exhaust heat exchanger must incorporate means to prevent 
blockage of the exhaust port after any internal heat exchanger failure.
    (i) For the purpose of compliance with Sec. 23.603, the failure of 
any part of the exhaust system will be considered to adversely affect 
safety.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13095, Aug. 13, 1969; Amdt. 23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 
23-43, 58 FR 18974, Apr. 9, 1993; Amdt. 23-51, 61 FR 5137, Feb. 9, 1996]



Sec. 23.1123  Exhaust system.

    (a) Each exhaust system must be fireproof and corrosion-resistant, 
and must have means to prevent failure due to expansion by operating 
temperatures.
    (b) Each exhaust system must be supported to withstand the vibration 
and inertia loads to which it may be subjected in operation.
    (c) Parts of the system connected to components between which 
relative motion could exist must have means for flexibility.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18974, Apr. 9, 1993]



Sec. 23.1125  Exhaust heat exchangers.

    For reciprocating engine powered airplanes the following apply:
    (a) Each exhaust heat exchanger must be constructed and installed to 
withstand the vibration, inertia, and other loads that it may be 
subjected to in normal operation. In addition--
    (1) Each exchanger must be suitable for continued operation at high 
temperatures and resistant to corrosion from exhaust gases;
    (2) There must be means for inspection of critical parts of each 
exchanger; and
    (3) Each exchanger must have cooling provisions wherever it is 
subject to contact with exhaust gases.
    (b) Each heat exchanger used for heating ventilating air must be 
constructed so that exhaust gases may not enter the ventilating air.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 
41 FR 55465, Dec. 20, 1976]

[[Page 291]]

                   Powerplant Controls and Accessories



Sec. 23.1141  Powerplant controls: General.

    (a) Powerplant controls must be located and arranged under Sec. 
23.777 and marked under Sec. 23.1555(a).
    (b) Each flexible control must be shown to be suitable for the 
particular application.
    (c) Each control must be able to maintain any necessary position 
without--
    (1) Constant attention by flight crew members; or
    (2) Tendency to creep due to control loads or vibration.
    (d) Each control must be able to withstand operating loads without 
failure or excessive deflection.
    (e) For turbine engine powered airplanes, no single failure or 
malfunction, or probable combination thereof, in any powerplant control 
system may cause the failure of any powerplant function necessary for 
safety.
    (f) The portion of each powerplant control located in the engine 
compartment that is required to be operated in the event of fire must be 
at least fire resistant.
    (g) Powerplant valve controls located in the cockpit must have--
    (1) For manual valves, positive stops or in the case of fuel valves 
suitable index provisions, in the open and closed position; and
    (2) For power-assisted valves, a means to indicate to the flight 
crew when the valve--
    (i) Is in the fully open or fully closed position; or
    (ii) Is moving between the fully open and fully closed position.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13095, Aug. 13, 1969; Amdt. 23-14, 38 FR 31823, Nov. 19, 1973; Amdt. 
23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 23-51, 61 FR 5137, Feb. 9, 
1996]



Sec. 23.1142  Auxiliary power unit controls.

    Means must be provided on the flight deck for the starting, 
stopping, monitoring, and emergency shutdown of each installed auxiliary 
power unit.

[Doc. No. 26344, 58 FR 18974, Apr. 9, 1993]



Sec. 23.1143  Engine controls.

    (a) There must be a separate power or thrust control for each engine 
and a separate control for each supercharger that requires a control.
    (b) Power, thrust, and supercharger controls must be arranged to 
allow--
    (1) Separate control of each engine and each supercharger; and
    (2) Simultaneous control of all engines and all superchargers.
    (c) Each power, thrust, or supercharger control must give a positive 
and immediate responsive means of controlling its engine or 
supercharger.
    (d) The power, thrust, or supercharger controls for each engine or 
supercharger must be independent of those for every other engine or 
supercharger.
    (e) For each fluid injection (other than fuel) system and its 
controls not provided and approved as part of the engine, the applicant 
must show that the flow of the injection fluid is adequately controlled.
    (f) If a power, thrust, or a fuel control (other than a mixture 
control) incorporates a fuel shutoff feature, the control must have a 
means to prevent the inadvertent movement of the control into the off 
position. The means must--
    (1) Have a positive lock or stop at the idle position; and
    (2) Require a separate and distinct operation to place the control 
in the shutoff position.
    (g) For reciprocating single-engine airplanes, each power or thrust 
control must be designed so that if the control separates at the engine 
fuel metering device, the airplane is capable of continued safe flight 
and landing.

[Amdt. 23-7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23-17, 41 
FR 55465, Dec. 20, 1976; Amdt. 23-29, 49 FR 6847, Feb. 23, 1984; Amdt. 
23-43, 58 FR 18974, Apr. 9, 1993; Amdt. 23-51, 61 FR 5137, Feb. 9, 1996]



Sec. 23.1145  Ignition switches.

    (a) Ignition switches must control and shut off each ignition 
circuit on each engine.
    (b) There must be means to quickly shut off all ignition on 
multiengine airplanes by the grouping of switches or by a master 
ignition control.

[[Page 292]]

    (c) Each group of ignition switches, except ignition switches for 
turbine engines for which continuous ignition is not required, and each 
master ignition control must have a means to prevent its inadvertent 
operation.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 23-43, 58 FR 
18974, Apr. 9, 1993]



Sec. 23.1147  Mixture controls.

    (a) If there are mixture controls, each engine must have a separate 
control, and each mixture control must have guards or must be shaped or 
arranged to prevent confusion by feel with other controls.
    (1) The controls must be grouped and arranged to allow--
    (i) Separate control of each engine; and
    (ii) Simultaneous control of all engines.
    (2) The controls must require a separate and distinct operation to 
move the control toward lean or shut-off position.
    (b) For reciprocating single-engine airplanes, each manual engine 
mixture control must be designed so that, if the control separates at 
the engine fuel metering device, the airplane is capable of continued 
safe flight and landing.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13096, Aug. 13, 1969; Amdt. 23-33, 51 FR 26657, July 24, 1986; Amdt. 
23-43, 58 FR 18974, Apr. 9, 1993]



Sec. 23.1149  Propeller speed and pitch controls.

    (a) If there are propeller speed or pitch controls, they must be 
grouped and arranged to allow--
    (1) Separate control of each propeller; and
    (2) Simultaneous control of all propellers.
    (b) The controls must allow ready synchronization of all propellers 
on multiengine airplanes.



Sec. 23.1153  Propeller feathering controls.

    If there are propeller feathering controls installed, it must be 
possible to feather each propeller separately. Each control must have a 
means to prevent inadvertent operation.

[Doc. No. 27804, 61 FR 5138, Feb. 9, 1996]



Sec. 23.1155  Turbine engine reverse thrust and propeller pitch
settings below the flight regime.

    For turbine engine installations, each control for reverse thrust 
and for propeller pitch settings below the flight regime must have means 
to prevent its inadvertent operation. The means must have a positive 
lock or stop at the flight idle position and must require a separate and 
distinct operation by the crew to displace the control from the flight 
regime (forward thrust regime for turbojet powered airplanes).

[Amdt. 23-7, 34 FR 13096, Aug. 13, 1969]



Sec. 23.1157  Carburetor air temperature controls.

    There must be a separate carburetor air temperature control for each 
engine.



Sec. 23.1163  Powerplant accessories.

    (a) Each engine mounted accessory must--
    (1) Be approved for mounting on the engine involved and use the 
provisions on the engines for mounting; or
    (2) Have torque limiting means on all accessory drives in order to 
prevent the torque limits established for those drives from being 
exceeded; and
    (3) In addition to paragraphs (a)(1) or (a)(2) of this section, be 
sealed to prevent contamination of the engine oil system and the 
accessory system.
    (b) Electrical equipment subject to arcing or sparking must be 
installed to minimize the probability of contact with any flammable 
fluids or vapors that might be present in a free state.
    (c) Each generator rated at or more than 6 kilowatts must be 
designed and installed to minimize the probability of a fire hazard in 
the event it malfunctions.
    (d) If the continued rotation of any accessory remotely driven by 
the engine is hazardous when malfunctioning occurs, a means to prevent 
rotation without interfering with the continued operation of the engine 
must be provided.

[[Page 293]]

    (e) Each accessory driven by a gearbox that is not approved as part 
of the powerplant driving the gearbox must--
    (1) Have torque limiting means to prevent the torque limits 
established for the affected drive from being exceeded;
    (2) Use the provisions on the gearbox for mounting; and
    (3) Be sealed to prevent contamination of the gearbox oil system and 
the accessory system.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31823, Nov. 19, 1973; Amdt. 23-29, 49 FR 6847, Feb. 23, 1984; 
Amdt. 23-34, 52 FR 1832, Jan. 15, 1987; Amdt. 23-42, 56 FR 354, Jan. 3, 
1991]



Sec. 23.1165  Engine ignition systems.

    (a) Each battery ignition system must be supplemented by a generator 
that is automatically available as an alternate source of electrical 
energy to allow continued engine operation if any battery becomes 
depleted.
    (b) The capacity of batteries and generators must be large enough to 
meet the simultaneous demands of the engine ignition system and the 
greatest demands of any electrical system components that draw from the 
same source.
    (c) The design of the engine ignition system must account for--
    (1) The condition of an inoperative generator;
    (2) The condition of a completely depleted battery with the 
generator running at its normal operating speed; and
    (3) The condition of a completely depleted battery with the 
generator operating at idling speed, if there is only one battery.
    (d) There must be means to warn appropriate crewmembers if 
malfunctioning of any part of the electrical system is causing the 
continuous discharge of any battery used for engine ignition.
    (e) Each turbine engine ignition system must be independent of any 
electrical circuit that is not used for assisting, controlling, or 
analyzing the operation of that system.
    (f) In addition, for commuter category airplanes, each turbine 
engine ignition system must be an essential electrical load.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 
41 FR 55465 Dec. 20, 1976; Amdt. 23-34, 52 FR 1833, Jan. 15, 1987; Amdt. 
23-62, 76 FR 75759, Dec. 2, 2011]

                       Powerplant Fire Protection



Sec. 23.1181  Designated fire zones; regions included.

    Designated fire zones are--
    (a) For reciprocating engines--
    (1) The power section;
    (2) The accessory section;
    (3) Any complete powerplant compartment in which there is no 
isolation between the power section and the accessory section.
    (b) For turbine engines--
    (1) The compressor and accessory sections;
    (2) The combustor, turbine and tailpipe sections that contain lines 
or components carrying flammable fluids or gases.
    (3) Any complete powerplant compartment in which there is no 
isolation between compressor, accessory, combustor, turbine, and 
tailpipe sections.
    (c) Any auxiliary power unit compartment; and
    (d) Any fuel-burning heater, and other combustion equipment 
installation described in Sec. 23.859.

[Doc. No. 26344, 58 FR 18975, Apr. 9, 1993, as amended by Amdt. 23-51, 
61 FR 5138, Feb. 9, 1996]



Sec. 23.1182  Nacelle areas behind firewalls.

    Components, lines, and fittings, except those subject to the 
provisions of Sec. 23.1351(e), located behind the engine-compartment 
firewall must be constructed of such materials and located at such 
distances from the firewall that they will not suffer damage sufficient 
to endanger the airplane if a portion of the engine side of the firewall 
is subjected to a flame temperature of not less than 2000 [deg]F for 15 
minutes.

[Amdt. 23-14, 38 FR 31816, Nov. 19, 1973]



Sec. 23.1183  Lines, fittings, and components.

    (a) Except as provided in paragraph (b) of this section, each 
component, line, and fitting carrying flammable

[[Page 294]]

fluids, gas, or air in any area subject to engine fire conditions must 
be at least fire resistant, except that flammable fluid tanks and 
supports which are part of and attached to the engine must be fireproof 
or be enclosed by a fireproof shield unless damage by fire to any non-
fireproof part will not cause leakage or spillage of flammable fluid. 
Components must be shielded or located so as to safeguard against the 
ignition of leaking flammable fluid. Flexible hose assemblies (hose and 
end fittings) must be shown to be suitable for the particular 
application. An integral oil sump of less than 25-quart capacity on a 
reciprocating engine need not be fireproof nor be enclosed by a 
fireproof shield.
    (b) Paragraph (a) of this section does not apply to--
    (1) Lines, fittings, and components which are already approved as 
part of a type certificated engine; and
    (2) Vent and drain lines, and their fittings, whose failure will not 
result in, or add to, a fire hazard.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-5, 32 
FR 6912, May 5, 1967; Amdt. 23-15, 39 FR 35460, Oct. 1, 1974; Amdt. 23-
29, 49 FR 6847, Feb. 23, 1984; Amdt. 23-51, 61 FR 5138, Feb. 9, 1996]



Sec. 23.1189  Shutoff means.

    (a) For each multiengine airplane the following apply:
    (1) Each engine installation must have means to shut off or 
otherwise prevent hazardous quantities of fuel, oil, deicing fluid, and 
other flammable liquids from flowing into, within, or through any engine 
compartment, except in lines, fittings, and components forming an 
integral part of an engine.
    (2) The closing of the fuel shutoff valve for any engine may not 
make any fuel unavailable to the remaining engines that would be 
available to those engines with that valve open.
    (3) Operation of any shutoff means may not interfere with the later 
emergency operation of other equipment such as propeller feathering 
devices.
    (4) Each shutoff must be outside of the engine compartment unless an 
equal degree of safety is provided with the shutoff inside the 
compartment.
    (5) Not more than one quart of flammable fluid may escape into the 
engine compartment after engine shutoff. For those installations where 
the flammable fluid that escapes after shutdown cannot be limited to one 
quart, it must be demonstrated that this greater amount can be safely 
contained or drained overboard.
    (6) There must be means to guard against inadvertent operation of 
each shutoff means, and to make it possible for the crew to reopen the 
shutoff means in flight after it has been closed.
    (b) Turbine engine installations need not have an engine oil system 
shutoff if--
    (1) The oil tank is integral with, or mounted on, the engine; and
    (2) All oil system components external to the engine are fireproof 
or located in areas not subject to engine fire conditions.
    (c) Power operated valves must have means to indicate to the flight 
crew when the valve has reached the selected position and must be 
designed so that the valve will not move from the selected position 
under vibration conditions likely to exist at the valve location.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13096, Aug. 13, 1969; Amdt. 23-14, 38 FR 31823, Nov. 19, 1973; Amdt. 
23-29, 49 FR 6847, Feb. 23, 1984; Amdt. 23-43, 58 FR 18975, Apr. 9, 
1993]



Sec. 23.1191  Firewalls.

    (a) Each engine, auxiliary power unit, fuel burning heater, and 
other combustion equipment, must be isolated from the rest of the 
airplane by firewalls, shrouds, or equivalent means.
    (b) Each firewall or shroud must be constructed so that no hazardous 
quantity of liquid, gas, or flame can pass from the compartment created 
by the firewall or shroud to other parts of the airplane.
    (c) Each opening in the firewall or shroud must be sealed with close 
fitting, fireproof grommets, bushings, or firewall fittings.
    (d) [Reserved]
    (e) Each firewall and shroud must be fireproof and protected against 
corrosion.
    (f) Compliance with the criteria for fireproof materials or 
components must be shown as follows:

[[Page 295]]

    (1) The flame to which the materials or components are subjected 
must be 2,000 150 [deg]F.
    (2) Sheet materials approximately 10 inches square must be subjected 
to the flame from a suitable burner.
    (3) The flame must be large enough to maintain the required test 
temperature over an area approximately five inches square.
    (g) Firewall materials and fittings must resist flame penetration 
for at least 15 minutes.
    (h) The following materials may be used in firewalls or shrouds 
without being tested as required by this section:
    (1) Stainless steel sheet, 0.015 inch thick.
    (2) Mild steel sheet (coated with aluminum or otherwise protected 
against corrosion) 0.018 inch thick.
    (3) Terne plate, 0.018 inch thick.
    (4) Monel metal, 0.018 inch thick.
    (5) Steel or copper base alloy firewall fittings.
    (6) Titanium sheet, 0.016 inch thick.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18975, Apr. 9, 1993; 58 FR 27060, May 6, 1993; Amdt. 23-51, 61 FR 
5138, Feb. 9, 1996]



Sec. 23.1192  Engine accessory compartment diaphragm.

    For aircooled radial engines, the engine power section and all 
portions of the exhaust sytem must be isolated from the engine accessory 
compartment by a diaphragm that meets the firewall requirements of Sec. 
23.1191.

[Amdt. 23-14, 38 FR 31823, Nov. 19, 1973]



Sec. 23.1193  Cowling and nacelle.

    (a) Each cowling must be constructed and supported so that it can 
resist any vibration, inertia, and air loads to which it may be 
subjected in operation.
    (b) There must be means for rapid and complete drainage of each part 
of the cowling in the normal ground and flight attitudes. Drain 
operation may be shown by test, analysis, or both, to ensure that under 
normal aerodynamic pressure distribution expected in service each drain 
will operate as designed. No drain may discharge where it will cause a 
fire hazard.
    (c) Cowling must be at least fire resistant.
    (d) Each part behind an opening in the engine compartment cowling 
must be at least fire resistant for a distance of at least 24 inches aft 
of the opening.
    (e) Each part of the cowling subjected to high temperatures due to 
its nearness to exhaust sytem ports or exhaust gas impingement, must be 
fire proof.
    (f) Each nacelle of a multiengine airplane with supercharged engines 
must be designed and constructed so that with the landing gear 
retracted, a fire in the engine compartment will not burn through a 
cowling or nacelle and enter a nacelle area other than the engine 
compartment.
    (g) In addition, for all airplanes with engine(s) embedded in the 
fuselage or in pylons on the aft fuselage, the airplane must be designed 
so that no fire originating in any engine compartment can enter, either 
through openings or by burn-through, any other region where it would 
create additional hazards.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 23-34, 52 FR 
1833, Jan. 15, 1987; 58 FR 18975, Apr. 9, 1993; Amdt. 23-62, 76 FR 
75759, Dec. 2, 2011]



Sec. 23.1195  Fire extinguishing systems.

    (a) For all airplanes with engine(s) embedded in the fuselage or in 
pylons on the aft fuselage, fire extinguishing systems must be installed 
and compliance shown with the following:
    (1) Except for combustor, turbine, and tailpipe sections of turbine-
engine installations that contain lines or components carrying flammable 
fluids or gases for which a fire originating in these sections is shown 
to be controllable, a fire extinguisher system must serve each engine 
compartment;
    (2) The fire extinguishing system, the quantity of the extinguishing 
agent, the rate of discharge, and the discharge distribution must be 
adequate to extinguish fires. An individual ``one shot'' system may be 
used, except for engine(s) embedded in the fuselage, where a ``two 
shot'' system is required.
    (3) The fire extinguishing system for a nacelle must be able to 
simultaneously protect each compartment of the nacelle for which 
protection is provided.

[[Page 296]]

    (b) If an auxiliary power unit is installed in any airplane 
certificated to this part, that auxiliary power unit compartment must be 
served by a fire extinguishing system meeting the requirements of 
paragraph (a)(2) of this section.

[Amdt. 23-34, 52 FR 1833, Jan. 15, 1987, as amended by Amdt. 23-43, 58 
FR 18975, Apr. 9, 1993; Amdt. 23-62, 76 FR 75759, Dec. 2, 2011]



Sec. 23.1197  Fire extinguishing agents.

    For all airplanes with engine(s) embedded in the fuselage or in 
pylons on the aft fuselage the following applies:
    (1) Be capable of extinguishing flames emanating from any burning of 
fluids or other combustible materials in the area protected by the fire 
extinguishing system; and
    (2) Have thermal stability over the temperature range likely to be 
experienced in the compartment in which they are stored.
    (b) If any toxic extinguishing agent is used, provisions must be 
made to prevent harmful concentrations of fluid or fluid vapors (from 
leakage during normal operation of the airplane or as a result of 
discharging the fire extinguisher on the ground or in flight) from 
entering any personnel compartment, even though a defect may exist in 
the extinguishing system. This must be shown by test except for built-in 
carbon dioxide fuselage compartment fire extinguishing systems for 
which--
    (1) Five pounds or less of carbon dioxide will be discharged, under 
established fire control procedures, into any fuselage compartment; or
    (2) Protective breathing equipment is available for each flight 
crewmember on flight deck duty.

[Amdt. 23-34, 52 FR 1833, Jan. 15, 1987, as amended by Amdt. 23-62, 76 
FR 75760, Dec. 2, 2011]



Sec. 23.1199  Extinguishing agent containers.

    For all airplanes with engine(s) embedded in the fuselage or in 
pylons on the aft fuselage the following applies:
    (a) Each extinguishing agent container must have a pressure relief 
to prevent bursting of the container by excessive internal pressures.
    (b) The discharge end of each discharge line from a pressure relief 
connection must be located so that discharge of the fire extinguishing 
agent would not damage the airplane. The line must also be located or 
protected to prevent clogging caused by ice or other foreign matter.
    (c) A means must be provided for each fire extinguishing agent 
container to indicate that the container has discharged or that the 
charging pressure is below the established minimum necessary for proper 
functioning.
    (d) The temperature of each container must be maintained, under 
intended operating conditions, to prevent the pressure in the container 
from--
    (1) Falling below that necessary to provide an adequate rate of 
discharge; or
    (2) Rising high enough to cause premature discharge.
    (e) If a pyrotechnic capsule is used to discharge the extinguishing 
agent, each container must be installed so that temperature conditions 
will not cause hazardous deterioration of the pyrotechnic capsule.

[Amdt. 23-34, 52 FR 1833, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; 
Amdt. 23-62, 76 FR 75760, Dec. 2, 2011]



Sec. 23.1201  Fire extinguishing systems materials.

    For all airplanes with engine(s) embedded in the fuselage or in 
pylons on the aft fuselage the following applies:
    (a) No material in any fire extinguishing system may react 
chemically with any extinguishing agent so as to create a hazard.
    (b) Each system component in an engine compartment must be 
fireproof.

[Amdt. 23-34, 52 FR 1833, Jan. 15, 1987; 52 FR 7262, Mar. 9, 1987; Amdt. 
23-62, 76 FR 75760, Dec. 2, 2011]



Sec. 23.1203  Fire detector system.

    (a) There must be means that ensure the prompt detection of a fire 
in--
    (1) An engine compartment of--
    (i) Multiengine turbine powered airplanes;
    (ii) Multiengine reciprocating engine powered airplanes 
incorporating turbochargers;
    (iii) Airplanes with engine(s) located where they are not readily 
visible from the cockpit; and

[[Page 297]]

    (iv) All commuter category airplanes.
    (2) The auxiliary power unit compartment of any airplane 
incorporating an auxiliary power unit.
    (b) Each fire detector must be constructed and installed to 
withstand the vibration, inertia, and other loads to which it may be 
subjected in operation.
    (c) No fire detector may be affected by any oil, water, other 
fluids, or fumes that might be present.
    (d) There must be means to allow the crew to check, in flight, the 
functioning of each fire detector electric circuit.
    (e) Wiring and other components of each fire detector system in a 
designated fire zone must be at least fire resistant.

[Amdt. 23-18, 42 FR 15042, Mar. 17, 1977, as amended by Amdt. 23-34, 52 
FR 1833, Jan. 15, 1987; Amdt. 23-43, 58 FR 18975, Apr. 9, 1993; Amdt. 
23-51, 61 FR 5138, Feb. 9, 1996]



                           Subpart F_Equipment

                                 General



Sec. 23.1301  Function and installation.

    Each item of installed equipment must--
    (a) Be of a kind and design appropriate to its intended function.
    (b) Be labeled as to its identification, function, or operating 
limitations, or any applicable combination of these factors; and
    (c) Be installed according to limitations specified for that 
equipment.

[Amdt. 23-20, 42 FR 36968, July 18, 1977, as amended by Amdt. 23-62, 76 
FR 75760, Dec. 2, 2011]



Sec. 23.1303  Flight and navigation instruments.

    The following are the minimum required flight and navigation 
instruments:
    (a) An airspeed indicator.
    (b) An altimeter.
    (c) A magnetic direction indicator.
    (d) For reciprocating engine-powered airplanes of more than 6,000 
pounds maximum weight and turbine engine powered airplanes, a free air 
temperature indicator or an air-temperature indicator which provides 
indications that are convertible to free-air.
    (e) A speed warning device for--
    (1) Turbine engine powered airplanes; and
    (2) Other airplanes for which Vmo/Mmo and Vd/Md are established 
under Sec. Sec. 23.335(b)(4) and 23.1505(c) if Vmo/Mmo is greater than 
0.8 Vd/Md.
    The speed warning device must give effective aural warning 
(differing distinctively from aural warnings used for other purposes) to 
the pilots whenever the speed exceeds Vmo plus 6 knots or Mmo+0.01. The 
upper limit of the production tolerance for the warning device may not 
exceed the prescribed warning speed. The lower limit of the warning 
device must be set to minimize nuisance warning;
    (f) When an attitude display is installed, the instrument design 
must not provide any means, accessible to the flightcrew, of adjusting 
the relative positions of the attitude reference symbol and the horizon 
line beyond that necessary for parallax correction.
    (g) In addition, for commuter category airplanes:
    (1) If airspeed limitations vary with altitude, the airspeed 
indicator must have a maximum allowable airspeed indicator showing the 
variation of VMO with altitude.
    (2) The altimeter must be a sensitive type.
    (3) Having a passenger seating configuration of 10 or more, 
excluding the pilot's seats and that are approved for IFR operations, a 
third attitude instrument must be provided that:
    (i) Is powered from a source independent of the electrical 
generating system;
    (ii) Continues reliable operation for a minimum of 30 minutes after 
total failure of the electrical generating system;
    (iii) Operates independently of any other attitude indicating 
system;
    (iv) Is operative without selection after total failure of the 
electrical generating system;
    (v) Is located on the instrument panel in a position acceptable to 
the Administrator that will make it plainly visible to and usable by any 
pilot at the pilot's station; and

[[Page 298]]

    (vi) Is appropriately lighted during all phases of operation.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 
41 FR 55465, Dec. 20, 1976; Amdt. 23-43, 58 FR 18975, Apr. 9, 1993; 
Amdt. 23-49, 61 FR 5168, Feb. 9, 1996; Amdt. 23-62, 76 FR 75760, Dec. 2, 
2011]



Sec. 23.1305  Powerplant instruments.

    The following are required powerplant instruments:
    (a) For all airplanes. (1) A fuel quantity indicator for each fuel 
tank, installed in accordance with Sec. 23.1337(b).
    (2) An oil pressure indicator for each engine.
    (3) An oil temperature indicator for each engine.
    (4) An oil quantity measuring device for each oil tank which meets 
the requirements of Sec. 23.1337(d).
    (5) A fire warning means for those airplanes required to comply with 
Sec. 23.1203.
    (b) For reciprocating engine-powered airplanes. In addition to the 
powerplant instruments required by paragraph (a) of this section, the 
following powerplant instruments are required:
    (1) An induction system air temperature indicator for each engine 
equipped with a preheater and having induction air temperature 
limitations that can be exceeded with preheat.
    (2) A tachometer indicator for each engine.
    (3) A cylinder head temperature indicator for--
    (i) Each air-cooled engine with cowl flaps;
    (ii) [Reserved]
    (iii) Each commuter category airplane.
    (4) For each pump-fed engine, a means:
    (i) That continuously indicates, to the pilot, the fuel pressure or 
fuel flow; or
    (ii) That continuously monitors the fuel system and warns the pilot 
of any fuel flow trend that could lead to engine failure.
    (5) A manifold pressure indicator for each altitude engine and for 
each engine with a controllable propeller.
    (6) For each turbocharger installation:
    (i) If limitations are established for either carburetor (or 
manifold) air inlet temperature or exhaust gas or turbocharger turbine 
inlet temperature, indicators must be furnished for each temperature for 
which the limitation is established unless it is shown that the 
limitation will not be exceeded in all intended operations.
    (ii) If its oil system is separate from the engine oil system, oil 
pressure and oil temperature indicators must be provided.
    (7) A coolant temperature indicator for each liquid-cooled engine.
    (c) For turbine engine-powered airplanes. In addition to the 
powerplant instruments required by paragraph (a) of this section, the 
following powerplant instruments are required:
    (1) A gas temperature indicator for each engine.
    (2) A fuel flowmeter indicator for each engine.
    (3) A fuel low pressure warning means for each engine.
    (4) A fuel low level warning means for any fuel tank that should not 
be depleted of fuel in normal operations.
    (5) A tachometer indicator (to indicate the speed of the rotors with 
established limiting speeds) for each engine.
    (6) An oil low pressure warning means for each engine.
    (7) An indicating means to indicate the functioning of the 
powerplant ice protection system for each engine.
    (8) For each engine, an indicating means for the fuel strainer or 
filter required by Sec. 23.997 to indicate the occurrence of 
contamination of the strainer or filter before it reaches the capacity 
established in accordance with Sec. 23.997(d).
    (9) For each engine, a warning means for the oil strainer or filter 
required by Sec. 23.1019, if it has no bypass, to warn the pilot of the 
occurrence of contamination of the strainer or filter screen before it 
reaches the capacity established in accordance with Sec. 23.1019(a)(5).
    (10) An indicating means to indicate the functioning of any heater 
used to prevent ice clogging of fuel system components.
    (d) For turbojet/turbofan engine-powered airplanes. In addition to 
the powerplant instruments required by paragraphs (a) and (c) of this 
section, the following powerplant instruments are required:

[[Page 299]]

    (1) For each engine, an indicator to indicate thrust or to indicate 
a parameter that can be related to thrust, including a free air 
temperature indicator if needed for this purpose.
    (2) For each engine, a position indicating means to indicate to the 
flight crew when the thrust reverser, if installed, is in the reverse 
thrust position.
    (e) For turbopropeller-powered airplanes. In addition to the 
powerplant instruments required by paragraphs (a) and (c) of this 
section, the following powerplant instruments are required:
    (1) A torque indicator for each engine.
    (2) A position indicating means to indicate to the flight crew when 
the propeller blade angle is below the flight low pitch position, for 
each propeller, unless it can be shown that such occurrence is highly 
improbable.

[Doc. No. 26344, 58 FR 18975, Apr. 9, 1993; 58 FR 27060, May 6, 1993; 
Amdt. 23-51, 61 FR 5138, Feb. 9, 1996; Amdt. 23-52, 61 FR 13644, Mar. 
27, 1996]



Sec. 23.1306  Electrical and electronic system lightning protection.

    (a) Each electrical and electronic system that performs a function, 
for which failure would prevent the continued safe flight and landing of 
the airplane, must be designed and installed so that--
    (1) The function is not adversely affected during and after the time 
the airplane is exposed to lightning; and
    (2) The system automatically recovers normal operation of that 
function in a timely manner after the airplane is exposed to lightning.
    (b) For airplanes approved for instrument flight rules operation, 
each electrical and electronic system that performs a function, for 
which failure would reduce the capability of the airplane or the ability 
of the flightcrew to respond to an adverse operating condition, must be 
designed and installed so that the function recovers normal operation in 
a timely manner after the airplane is exposed to lightning.

[Doc. No. FAA-2010-0224; 76 FR 33135, June 8, 2011]



Sec. 23.1307  Miscellaneous equipment.

    The equipment necessary for an airplane to operate at the maximum 
operating altitude and in the kinds of operation and meteorological 
conditions for which certification is requested and is approved in 
accordance with Sec. 23.1559 must be included in the type design.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-23, 43 FR 50593, Oct. 30, 1978; Amdt. 23-43, 58 FR 
18976, Apr. 9, 1993; Amdt. 23-49, 61 FR 5168, Feb. 9, 1996]



Sec. 23.1308  High-intensity Radiated Fields (HIRF) Protection.

    (a) Except as provided in paragraph (d) of this section, each 
electrical and electronic system that performs a function whose failure 
would prevent the continued safe flight and landing of the airplane must 
be designed and installed so that--
    (1) The function is not adversely affected during and after the time 
the airplane is exposed to HIRF environment I, as described in appendix 
J to this part;
    (2) The system automatically recovers normal operation of that 
function, in a timely manner, after the airplane is exposed to HIRF 
environment I, as described in appendix J to this part, unless the 
system's recovery conflicts with other operational or functional 
requirements of the system; and
    (3) The system is not adversely affected during and after the time 
the airplane is exposed to HIRF environment II, as described in appendix 
J to this part.
    (b) Each electrical and electronic system that performs a function 
whose failure would significantly reduce the capability of the airplane 
or the ability of the flightcrew to respond to an adverse operating 
condition must be designed and installed so the system is not adversely 
affected when the equipment providing the function is exposed to 
equipment HIRF test level 1 or 2, as described in appendix J to this 
part.
    (c) Each electrical and electronic system that performs a function 
whose failure would reduce the capability of the airplane or the ability 
of the flightcrew to respond to an adverse operating condition must be 
designed and installed so the system is not adversely

[[Page 300]]

affected when the equipment providing the function is exposed to 
equipment HIRF test level 3, as described in appendix J to this part.
    (d) Before December 1, 2012, an electrical or electronic system that 
performs a function whose failure would prevent the continued safe 
flight and landing of an airplane may be designed and installed without 
meeting the provisions of paragraph (a) provided--
    (1) The system has previously been shown to comply with special 
conditions for HIRF, prescribed under Sec. 21.16, issued before 
December 1, 2007;
    (2) The HIRF immunity characteristics of the system have not changed 
since compliance with the special conditions was demonstrated; and
    (3) The data used to demonstrate compliance with the special 
conditions is provided.

[Doc. No. FAA-2006-23657, 72 FR 44024, Aug. 6, 2007]



Sec. 23.1309  Equipment, systems, and installations.

    The requirements of this section, except as identified in paragraphs 
(a) through (d), are applicable, in addition to specific design 
requirements of part 23, to any equipment or system as installed in the 
airplane. This section is a regulation of general requirements and does 
not supersede any requirements contained in another section of part 23.
    (a) The airplane equipment and systems must be designed and 
installed so that:
    (1) Those required for type certification or by operating rules 
perform as intended under the airplane operating and environmental 
conditions, including the indirect effects of lightning strikes.
    (2) Any equipment and system does not adversely affect the safety of 
the airplane or its occupants, or the proper functioning of those 
covered by paragraph (a)(1) of this section.
    (b) Minor, major, hazardous, or catastrophic failure condition(s), 
which occur during Type Inspection Authorization or FAA flight-
certification testing, must have root cause analysis and corrective 
action.
    (c) The airplane systems and associated components considered 
separately and in relation to other systems, must be designed and 
installed so that:
    (1) Each catastrophic failure condition is extremely improbable and 
does not result from a single failure;
    (2) Each hazardous failure condition is extremely remote; and
    (3) Each major failure condition is remote.
    (d) Information concerning an unsafe system operating condition must 
be provided in a timely manner to the crew to enable them to take 
appropriate corrective action. An appropriate alert must be provided if 
immediate pilot awareness and immediate or subsequent corrective action 
is required. Systems and controls, including indications and 
annunciations, must be designed to minimize crew errors which could 
create additional hazards.

[Doc. No. FAA-2009-0738, 76 FR 75760, Dec. 2, 2011]



Sec. 23.1310  Power source capacity and distribution.

    (a) Each installation whose functioning is required for type 
certification or under operating rules and that requires a power supply 
is an ``essential load'' on the power supply. The power sources and the 
system must be able to supply the following power loads in probable 
operating combinations and for probable durations:
    (1) Loads connected to the system with the system functioning 
normally.
    (2) Essential loads, after failure of any one prime mover, power 
converter, or energy storage device.
    (3) Essential loads after failure of--
    (i) Any one engine on two-engine airplanes; and
    (ii) Any two engines on airplanes with three or more engines.
    (4) Essential loads for which an alternate source of power is 
required, after any failure or malfunction in any one power supply 
system, distribution system, or other utilization system.
    (b) In determining compliance with paragraphs (a)(2) and (3) of this 
section, the power loads may be assumed to be reduced under a monitoring 
procedure consistent with safety in the kinds of operation authorized. 
Loads not required in controlled flight need not be

[[Page 301]]

considered for the two-engine-inoperative condition on airplanes with 
three or more engines.

[Doc. No. FAA-2009-0738, 76 FR 75760, Dec. 2, 2011]

                        Instruments: Installation



Sec. 23.1311  Electronic display instrument systems.

    (a) Electronic display indicators, including those with features 
that make isolation and independence between powerplant instrument 
systems impractical, must:
    (1) Meet the arrangement and visibility requirements of Sec. 
23.1321.
    (2) Be easily legible under all lighting conditions encountered in 
the cockpit, including direct sunlight, considering the expected 
electronic display brightness level at the end of an electronic display 
indictor's useful life. Specific limitations on display system useful 
life must be contained in the Instructions for Continued Airworthiness 
required by Sec. 23.1529.
    (3) Not inhibit the primary display of attitude, airspeed, altitude, 
or powerplant parameters needed by any pilot to set power within 
established limitations, in any normal mode of operation.
    (4) Not inhibit the primary display of engine parameters needed by 
any pilot to properly set or monitor powerplant limitations during the 
engine starting mode of operation.
    (5) For certification for Instrument Flight Rules (IFR) operations, 
have an independent magnetic direction indicator and either an 
independent secondary mechanical altimeter, airspeed indicator, and 
attitude instrument or an electronic display parameters for the 
altitude, airspeed, and attitude that are independent from the 
airplane's primary electrical power system. These secondary instruments 
may be installed in panel positions that are displaced from the primary 
positions specified by Sec. 23.1321(d), but must be located where they 
meet the pilot's visibility requirements of Sec. 23.1321(a).
    (6) Incorporate sensory cues that provide a quick glance sense of 
rate and, where appropriate, trend information to the parameter being 
displayed to the pilot.
    (7) Incorporate equivalent visual displays of the instrument 
markings required by Sec. Sec. 23.1541 through 23.1553, or visual 
displays that alert the pilot to abnormal operational values or 
approaches to established limitation values, for each parameter required 
to be displayed by this part.
    (b) The electronic display indicators, including their systems and 
installations, and considering other airplane systems, must be designed 
so that one display of information essential for continued safe flight 
and landing will be available within one second to the crew by a single 
pilot action or by automatic means for continued safe operation, after 
any single failure or probable combination of failures.
    (c) As used in this section, ``instrument'' includes devices that 
are physically contained in one unit, and devices that are composed of 
two or more physically separate units or components connected together 
(such as a remote indicating gyroscopic direction indicator that 
includes a magnetic sensing element, a gyroscopic unit, an amplifier, 
and an indicator connected together). As used in this section, 
``primary'' display refers to the display of a parameter that is located 
in the instrument panel such that the pilot looks at it first when 
wanting to view that parameter.

[Doc. No. 27806, 61 FR 5168, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75760, Dec. 2, 2011]



Sec. 23.1321  Arrangement and visibility.

    (a) Each flight, navigation, and powerplant instrument for use by 
any required pilot during takeoff, initial climb, final approach, and 
landing must be located so that any pilot seated at the controls can 
monitor the airplane's flight path and these instruments with minimum 
head and eye movement. The powerplant instruments for these flight 
conditions are those needed to set power within powerplant limitations.
    (b) For each multiengine airplane, identical powerplant instruments 
must be located so as to prevent confusion as to which engine each 
instrument relates.

[[Page 302]]

    (c) Instrument panel vibration may not damage, or impair the 
accuracy of, any instrument.
    (d) For each airplane, the flight instruments required by Sec. 
23.1303, and, as applicable, by the operating rules of this chapter, 
must be grouped on the instrument panel and centered as nearly as 
practicable about the vertical plane of each required pilot's forward 
vision. In addition:
    (1) The instrument that most effectively indicates the attitude must 
be on the panel in the top center position;
    (2) The instrument that most effectively indicates airspeed must be 
adjacent to and directly to the left of the instrument in the top center 
position;
    (3) The instrument that most effectively indicates altitude must be 
adjacent to and directly to the right of the instrument in the top 
center position;
    (4) The instrument that most effectively indicates direction of 
flight, other than the magnetic direction indicator required by Sec. 
23.1303(c), must be adjacent to and directly below the instrument in the 
top center position; and
    (5) Electronic display indicators may be used for compliance with 
paragraphs (d)(1) through (d)(4) of this section when such displays 
comply with requirements in Sec. 23.1311.
    (e) If a visual indicator is provided to indicate malfunction of an 
instrument, it must be effective under all probable cockpit lighting 
conditions.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31824, Nov. 19, 1973; Amdt. 23-20, 42 FR 36968, July 18, 1977; 
Amdt. 23-41, 55 FR 43310, Oct. 26, 1990; 55 FR 46888, Nov. 7, 1990; 
Amdt. 23-49, 61 FR 5168, Feb. 9, 1996]



Sec. 23.1322  Warning, caution, and advisory lights.

    If warning, caution, or advisory lights are installed in the 
cockpit, they must, unless otherwise approved by the Administrator, be--
    (a) Red, for warning lights (lights indicating a hazard which may 
require immediate corrective action);
    (b) Amber, for caution lights (lights indicating the possible need 
for future corrective action);
    (c) Green, for safe operation lights; and
    (d) Any other color, including white, for lights not described in 
paragraphs (a) through (c) of this section, provided the color differs 
sufficiently from the colors prescribed in paragraphs (a) through (c) of 
this section to avoid possible confusion.
    (e) Effective under all probable cockpit lighting conditions.

[Amdt. 23-17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23-43, 58 
FR 18976, Apr. 9, 1993]



Sec. 23.1323  Airspeed indicating system.

    (a) Each airspeed indicating instrument must be calibrated to 
indicate true airspeed (at sea level with a standard atmosphere) with a 
minimum practicable instrument calibration error when the corresponding 
pitot and static pressures are applied.
    (b) Each airspeed system must be calibrated in flight to determine 
the system error. The system error, including position error, but 
excluding the airspeed indicator instrument calibration error, may not 
exceed three percent of the calibrated airspeed or five knots, whichever 
is greater, throughout the following speed ranges:
    (1) 1.3 VS1 to VMO/MMO or 
VNE, whichever is appropriate with flaps retracted.
    (2) 1.3 VS1 to VFE with flaps extended.
    (c) The design and installation of each airspeed indicating system 
must provide positive drainage of moisture from the pitot static 
plumbing.
    (d) If certification for instrument flight rules or flight in icing 
conditions is requested, each airspeed system must have a heated pitot 
tube or an equivalent means of preventing malfunction due to icing.
    (e) In addition, for normal, utility, and acrobatic category 
multiengine jets of more than 6,000 pounds maximum weight and commuter 
category airplanes, each system must be calibrated to determine the 
system error during the accelerate-takeoff ground run. The ground run 
calibration must be determined--
    (1) From 0.8 of the minimum value of V1 to the maximum 
value of V2, considering the approved ranges of altitude and 
weight; and

[[Page 303]]

    (2) The ground run calibration must be determined assuming an engine 
failure at the minimum value of V1.
    (f) For commuter category airplanes, where duplicate airspeed 
indicators are required, their respective pitot tubes must be far enough 
apart to avoid damage to both tubes in a collision with a bird.

[Amdt. 23-20, 42 FR 36968, July 18, 1977, as amended by Amdt. 23-34, 52 
FR 1834, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23-42, 56 FR 
354, Jan. 3, 1991; Amdt. 23-49, 61 FR 5168, Feb. 9, 1996; Amdt. 23-62, 
76 FR 75761, Dec. 2, 2011]



Sec. 23.1325  Static pressure system.

    (a) Each instrument provided with static pressure case connections 
must be so vented that the influence of airplane speed, the opening and 
closing of windows, airflow variations, moisture, or other foreign 
matter will least affect the accuracy of the instruments except as noted 
in paragraph (b)(3) of this section.
    (b) If a static pressure system is necessary for the functioning of 
instruments, systems, or devices, it must comply with the provisions of 
paragraphs (b)(1) through (3) of this section.
    (1) The design and installation of a static pressure system must be 
such that--
    (i) Positive drainage of moisture is provided;
    (ii) Chafing of the tubing, and excessive distortion or restriction 
at bends in the tubing, is avoided; and
    (iii) The materials used are durable, suitable for the purpose 
intended, and protected against corrosion.
    (2) A proof test must be conducted to demonstrate the integrity of 
the static pressure system in the following manner:
    (i) Unpressurized airplanes. Evacuate the static pressure system to 
a pressure differential of approximately 1 inch of mercury or to a 
reading on the altimeter, 1,000 feet above the aircraft elevation at the 
time of the test. Without additional pumping for a period of 1 minute, 
the loss of indicated altitude must not exceed 100 feet on the 
altimeter.
    (ii) Pressurized airplanes. Evacuate the static pressure system 
until a pressure differential equivalent to the maximum cabin pressure 
differential for which the airplane is type certificated is achieved. 
Without additional pumping for a period of 1 minute, the loss of 
indicated altitude must not exceed 2 percent of the equivalent altitude 
of the maximum cabin differential pressure or 100 feet, whichever is 
greater.
    (3) If a static pressure system is provided for any instrument, 
device, or system required by the operating rules of this chapter, each 
static pressure port must be designed or located in such a manner that 
the correlation between air pressure in the static pressure system and 
true ambient atmospheric static pressure is not altered when the 
airplane encounters icing conditions. An antiicing means or an alternate 
source of static pressure may be used in showing compliance with this 
requirement. If the reading of the altimeter, when on the alternate 
static pressure system differs from the reading of the altimeter when on 
the primary static system by more than 50 feet, a correction card must 
be provided for the alternate static system.
    (c) Except as provided in paragraph (d) of this section, if the 
static pressure system incorporates both a primary and an alternate 
static pressure source, the means for selecting one or the other source 
must be designed so that--
    (1) When either source is selected, the other is blocked off; and
    (2) Both sources cannot be blocked off simultaneously.
    (d) For unpressurized airplanes, paragraph (c)(1) of this section 
does not apply if it can be demonstrated that the static pressure system 
calibration, when either static pressure source is selected, is not 
changed by the other static pressure source being open or blocked.
    (e) Each static pressure system must be calibrated in flight to 
determine the system error. The system error, in indicated pressure 
altitude, at sea-level, with a standard atmosphere, excluding instrument 
calibration error, may not exceed 30 feet per 100 
knot speed for the appropriate configuration in the speed range between 
1.3 VS0 with flaps extended, and 1.8 VS1 with 
flaps retracted. However, the error need not be less than 30 feet.

[[Page 304]]

    (f) [Reserved]
    (g) For airplanes prohibited from flight in instrument 
meteorological or icing conditions, in accordance with Sec. 23.1559(b) 
of this part, paragraph (b)(3) of this section does not apply.

[Amdt. 23-1, 30 FR 8261, June 29, 1965, as amended by Amdt. 23-6, 32 FR 
7586, May 24, 1967; 32 FR 13505, Sept. 27, 1967; 32 FR 13714, Sept. 30, 
1967; Amdt. 23-20, 42 FR 36968, July 18, 1977; Amdt. 23-34, 52 FR 1834, 
Jan. 15, 1987; Amdt. 23-42, 56 FR 354, Jan. 3, 1991; Amdt. 23-49, 61 FR 
5169, Feb. 9, 1996; Amdt. 23-50, 61 FR 5192, Feb. 9, 1996]



Sec. 23.1326  Pitot heat indication systems.

    If a flight instrument pitot heating system is installed to meet the 
requirements specified in Sec. 23.1323(d), an indication system must be 
provided to indicate to the flight crew when that pitot heating system 
is not operating. The indication system must comply with the following 
requirements:
    (a) The indication provided must incorporate an amber light that is 
in clear view of a flightcrew member.
    (b) The indication provided must be designed to alert the flight 
crew if either of the following conditions exist:
    (1) The pitot heating system is switched ``off.''
    (2) The pitot heating system is switched ``on'' and any pitot tube 
heating element is inoperative.

[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]



Sec. 23.1327  Magnetic direction indicator.

    (a) Except as provided in paragraph (b) of this section--
    (1) Each magnetic direction indicator must be installed so that its 
accuracy is not excessively affected by the airplane's vibration or 
magnetic fields; and
    (2) The compensated installation may not have a deviation in level 
flight, greater than ten degrees on any heading.
    (b) A magnetic nonstabilized direction indicator may deviate more 
than ten degrees due to the operation of electrically powered systems 
such as electrically heated windshields if either a magnetic stabilized 
direction indicator, which does not have a deviation in level flight 
greater than ten degrees on any heading, or a gyroscopic direction 
indicator, is installed. Deviations of a magnetic nonstabilized 
direction indicator of more than 10 degrees must be placarded in 
accordance with Sec. 23.1547(e).

[Amdt. 23-20, 42 FR 36969, July 18, 1977]



Sec. 23.1329  Automatic pilot system.

    If an automatic pilot system is installed, it must meet the 
following:
    (a) Each system must be designed so that the automatic pilot can--
    (1) Be quickly and positively disengaged by the pilots to prevent it 
from interfering with their control of the airplane; or
    (2) Be sufficiently overpowered by one pilot to let him control the 
airplane.
    (b) If the provisions of paragraph (a)(1) of this section are 
applied, the quick release (emergency) control must be located on the 
control wheel (both control wheels if the airplane can be operated from 
either pilot seat) on the side opposite the throttles, or on the stick 
control, (both stick controls, if the airplane can be operated from 
either pilot seat) such that it can be operated without moving the hand 
from its normal position on the control.
    (c) Unless there is automatic synchronization, each system must have 
a means to readily indicate to the pilot the alignment of the actuating 
device in relation to the control system it operates.
    (d) Each manually operated control for the system operation must be 
readily accessible to the pilot. Each control must operate in the same 
plane and sense of motion as specified in Sec. 23.779 for cockpit 
controls. The direction of motion must be plainly indicated on or near 
each control.
    (e) Each system must be designed and adjusted so that, within the 
range of adjustment available to the pilot, it cannot produce hazardous 
loads on the airplane or create hazardous deviations in the flight path, 
under any flight condition appropriate to its use, either during normal 
operation or in the event of a malfunction, assuming that corrective 
action begins within a reasonable period of time.
    (f) Each system must be designed so that a single malfunction will 
not produce a hardover signal in more than one control axis. If the 
automatic pilot

[[Page 305]]

integrates signals from auxiliary controls or furnishes signals for 
operation of other equipment, positive interlocks and sequencing of 
engagement to prevent improper operation are required.
    (g) There must be protection against adverse interaction of 
integrated components, resulting from a malfunction.
    (h) If the automatic pilot system can be coupled to airborne 
navigation equipment, means must be provided to indicate to the flight 
crew the current mode of operation. Selector switch position is not 
acceptable as a means of indication.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-23, 43 FR 50593, Oct. 30, 1978; Amdt. 23-43, 58 FR 
18976, Apr. 9, 1993; Amdt. 23-49, 61 FR 5169, Feb. 9, 1996]



Sec. 23.1331  Instruments using a power source.

    For each instrument that uses a power source, the following apply:
    (a) Each instrument must have an integral visual power annunciator 
or separate power indicator to indicate when power is not adequate to 
sustain proper instrument performance. If a separate indicator is used, 
it must be located so that the pilot using the instruments can monitor 
the indicator with minimum head and eye movement. The power must be 
sensed at or near the point where it enters the instrument. For electric 
and vacuum/pressure instruments, the power is considered to be adequate 
when the voltage or the vacuum/pressure, respectively, is within 
approved limits.
    (b) The installation and power supply systems must be designed so 
that--
    (1) The failure of one instrument will not interfere with the proper 
supply of energy to the remaining instrument; and
    (2) The failure of the energy supply from one source will not 
interfere with the proper supply of energy from any other source.
    (c) For certification for Instrument Flight Rules (IFR) operations 
and for the heading, altitude, airspeed, and attitude, there must be at 
least:
    (1) Two independent sources of power (not driven by the same engine 
on multiengine airplanes), and a manual or an automatic means to select 
each power source; or
    (2) A separate display of parameters for heading, altitude, 
airspeed, and attitude that has a power source independent from the 
airplane's primary electrical power system.

[Doc. No. 26344, 58 FR 18976, Apr. 9, 1993, as amended by Amdt. 23-62, 
76 FR 75761, Dec. 2, 2011]



Sec. 23.1335  Flight director systems.

    If a flight director system is installed, means must be provided to 
indicate to the flight crew its current mode of operation. Selector 
switch position is not acceptable as a means of indication.

[Amdt. 23-20, 42 FR 36969, July 18, 1977]



Sec. 23.1337  Powerplant instruments installation.

    (a) Instruments and instrument lines. (1) Each powerplant and 
auxiliary power unit instrument line must meet the requirements of Sec. 
23.993.
    (2) Each line carrying flammable fluids under pressure must--
    (i) Have restricting orifices or other safety devices at the source 
of pressure to prevent the escape of excessive fluid if the line fails; 
and
    (ii) Be installed and located so that the escape of fluids would not 
create a hazard.
    (3) Each powerplant and auxiliary power unit instrument that 
utilizes flammable fluids must be installed and located so that the 
escape of fluid would not create a hazard.
    (b) Fuel quantity indication. There must be a means to indicate to 
the flightcrew members the quantity of usable fuel in each tank during 
flight. An indicator calibrated in appropriate units and clearly marked 
to indicate those units must be used. In addition:
    (1) Each fuel quantity indicator must be calibrated to read ``zero'' 
during level flight when the quantity of fuel remaining in the tank is 
equal to the unusable fuel supply determined under Sec. 23.959(a);
    (2) Each exposed sight gauge used as a fuel quantity indicator must 
be protected against damage;
    (3) Each sight gauge that forms a trap in which water can collect 
and

[[Page 306]]

freeze must have means to allow drainage on the ground;
    (4) There must be a means to indicate the amount of usable fuel in 
each tank when the airplane is on the ground (such as by a stick gauge);
    (5) Tanks with interconnected outlets and airspaces may be 
considered as one tank and need not have separate indicators; and
    (6) No fuel quantity indicator is required for an auxiliary tank 
that is used only to transfer fuel to other tanks if the relative size 
of the tank, the rate of fuel transfer, and operating instructions are 
adequate to--
    (i) Guard against overflow; and
    (ii) Give the flight crewmembers prompt warning if transfer is not 
proceeding as planned.
    (c) Fuel flowmeter system. If a fuel flowmeter system is installed, 
each metering component must have a means to by-pass the fuel supply if 
malfunctioning of that component severely restricts fuel flow.
    (d) Oil quantity indicator. There must be a means to indicate the 
quantity of oil in each tank--
    (1) On the ground (such as by a stick gauge); and
    (2) In flight, to the flight crew members, if there is an oil 
transfer system or a reserve oil supply system.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13096, Aug. 13, 1969; Amdt. 23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 
23-43, 58 FR 18976, Apr. 9, 1993; Amdt. 23-51, 61 FR 5138, Feb. 9, 1996; 
Amdt. 23-49, 61 FR 5169, Feb. 9, 1996]

                    Electrical Systems and Equipment



Sec. 23.1351  General.

    (a) Electrical system capacity. Each electrical system must be 
adequate for the intended use. In addition--
    (1) Electric power sources, their transmission cables, and their 
associated control and protective devices, must be able to furnish the 
required power at the proper voltage to each load circuit essential for 
safe operation; and
    (2) Compliance with paragraph (a)(1) of this section must be shown 
as follows--
    (i) For normal, utility, and acrobatic category airplanes, by an 
electrical load analysis or by electrical measurements that account for 
the electrical loads applied to the electrical system in probable 
combinations and for probable durations; and
    (ii) For commuter category airplanes, by an electrical load analysis 
that accounts for the electrical loads applied to the electrical system 
in probable combinations and for probable durations.
    (b) Function. For each electrical system, the following apply:
    (1) Each system, when installed, must be--
    (i) Free from hazards in itself, in its method of operation, and in 
its effects on other parts of the airplane;
    (ii) Protected from fuel, oil, water, other detrimental substances, 
and mechanical damage; and
    (iii) So designed that the risk of electrical shock to crew, 
passengers, and ground personnel is reduced to a minimum.
    (2) Electric power sources must function properly when connected in 
combination or independently.
    (3) No failure or malfunction of any electric power source may 
impair the ability of any remaining source to supply load circuits 
essential for safe operation.
    (4) In addition, for commuter category airplanes, the following 
apply:
    (i) Each system must be designed so that essential load circuits can 
be supplied in the event of reasonably probable faults or open circuits 
including faults in heavy current carrying cables;
    (ii) A means must be accessible in flight to the flight crewmembers 
for the individual and collective disconnection of the electrical power 
sources from the system;
    (iii) The system must be designed so that voltage and frequency, if 
applicable, at the terminals of all essential load equipment can be 
maintained within the limits for which the equipment is designed during 
any probable operating conditions;
    (iv) If two independent sources of electrical power for particular 
equipment or systems are required, their electrical energy supply must 
be ensured by means such as duplicate electrical equipment, throwover 
switching, or multichannel or loop circuits separately routed; and

[[Page 307]]

    (v) For the purpose of complying with paragraph (b)(5) of this 
section, the distribution system includes the distribution busses, their 
associated feeders, and each control and protective device.
    (c) Generating system. There must be at least one generator/
alternator if the electrical system supplies power to load circuits 
essential for safe operation. In addition--
    (1) Each generator/alternator must be able to deliver its continuous 
rated power, or such power as is limited by its regulation system.
    (2) Generator/alternator voltage control equipment must be able to 
dependably regulate the generator/alternator output within rated limits.
    (3) Automatic means must be provided to prevent damage to any 
generator/alternator and adverse effects on the airplane electrical 
system due to reverse current. A means must also be provided to 
disconnect each generator/alternator from the battery and other 
generators/alternators.
    (4) There must be a means to give immediate warning to the flight 
crew of a failure of any generator/alternator.
    (5) Each generator/alternator must have an overvoltage control 
designed and installed to prevent damage to the electrical system, or to 
equipment supplied by the electrical system that could result if that 
generator/alternator were to develop an overvoltage condition.
    (d) Instruments. A means must exist to indicate to appropriate 
flight crewmembers the electric power system quantities essential for 
safe operation.
    (1) For normal, utility, and acrobatic category airplanes with 
direct current systems, an ammeter that can be switched into each 
generator feeder may be used and, if only one generator exists, the 
ammeter may be in the battery feeder.
    (2) For commuter category airplanes, the essential electric power 
system quantities include the voltage and current supplied by each 
generator.
    (e) Fire resistance. Electrical equipment must be so designed and 
installed that in the event of a fire in the engine compartment, during 
which the surface of the firewall adjacent to the fire is heated to 
2,000 [deg]F for 5 minutes or to a lesser temperature substantiated by 
the applicant, the equipment essential to continued safe operation and 
located behind the firewall will function satisfactorily and will not 
create an additional fire hazard.
    (f) External power. If provisions are made for connecting external 
power to the airplane, and that external power can be electrically 
connected to equipment other than that used for engine starting, means 
must be provided to ensure that no external power supply having a 
reverse polarity, or a reverse phase sequence, can supply power to the 
airplane's electrical system. The external power connection must be 
located so that its use will not result in a hazard to the airplane or 
ground personnel.
    (g) It must be shown by analysis, tests, or both, that the airplane 
can be operated safely in VFR conditions, for a period of not less than 
five minutes, with the normal electrical power (electrical power sources 
excluding the battery and any other standby electrical sources) 
inoperative, with critical type fuel (from the standpoint of flameout 
and restart capability), and with the airplane initially at the maximum 
certificated altitude. Parts of the electrical system may remain on if--
    (1) A single malfunction, including a wire bundle or junction box 
fire, cannot result in loss of the part turned off and the part turned 
on; and
    (2) The parts turned on are electrically and mechanically isolated 
from the parts turned off.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13096, Aug. 13, 1969; Amdt. 23-14, 38 FR 31824, Nov. 19, 1973; Amdt. 
23-17, 41 FR 55465, Dec. 20, 1976; Amdt. 23-20, 42 FR 36969, July 18, 
1977; Amdt. 23-34, 52 FR 1834, Jan. 15, 1987; 52 FR 34745, Sept. 14, 
1987; Amdt. 23-43, 58 FR 18976, Apr. 9, 1993; Amdt. 23-49, 61 FR 5169, 
Feb. 9, 1996]



Sec. 23.1353  Storage battery design and installation.

    (a) Each storage battery must be designed and installed as 
prescribed in this section.
    (b) Safe cell temperatures and pressures must be maintained during 
any probable charging and discharging condition. No uncontrolled 
increase in cell

[[Page 308]]

temperature may result when the battery is recharged (after previous 
complete discharge)--
    (1) At maximum regulated voltage or power;
    (2) During a flight of maximum duration; and
    (3) Under the most adverse cooling condition likely to occur in 
service.
    (c) Compliance with paragraph (b) of this section must be shown by 
tests unless experience with similar batteries and installations has 
shown that maintaining safe cell temperatures and pressures presents no 
problem.
    (d) No explosive or toxic gases emitted by any battery in normal 
operation, or as the result of any probable malfunction in the charging 
system or battery installation, may accumulate in hazardous quantities 
within the airplane.
    (e) No corrosive fluids or gases that may escape from the battery 
may damage surrounding structures or adjacent essential equipment.
    (f) Each nickel cadmium battery installation capable of being used 
to start an engine or auxiliary power unit must have provisions to 
prevent any hazardous effect on structure or essential systems that may 
be caused by the maximum amount of heat the battery can generate during 
a short circuit of the battery or of its individual cells.
    (g) Nickel cadmium battery installations capable of being used to 
start an engine or auxiliary power unit must have--
    (1) A system to control the charging rate of the battery 
automatically so as to prevent battery overheating;
    (2) A battery temperature sensing and over-temperature warning 
system with a means for disconnecting the battery from its charging 
source in the event of an over-temperature condition; or
    (3) A battery failure sensing and warning system with a means for 
disconnecting the battery from its charging source in the event of 
battery failure.
    (h)(1) In the event of a complete loss of the primary electrical 
power generating system, the battery must be capable of providing 
electrical power to those loads that are essential to continued safe 
flight and landing for:
    (i) At least 30 minutes for airplanes that are certificated with a 
maximum altitude of 25,000 feet or less; and
    (ii) At least 60 minutes for airplanes that are certificated with a 
maximum altitude over 25,000 feet.
    (2) The time period includes the time to recognize the loss of 
generated power and to take appropriate load shedding action.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-20, 42 FR 36969, July 18, 1977; Amdt. 23-21, 43 FR 
2319, Jan. 16, 1978; Amdt. 23-49, 61 FR 5169, Feb. 9, 1996; Amdt. 23-62, 
76 FR 75761, Dec. 2, 2011]



Sec. 23.1357  Circuit protective devices.

    (a) Protective devices, such as fuses or circuit breakers, must be 
installed in all electrical circuits other than--
    (1) Main circuits of starter motors used during starting only; and
    (2) Circuits in which no hazard is presented by their omission.
    (b) A protective device for a circuit essential to flight safety may 
not be used to protect any other circuit.
    (c) Each resettable circuit protective device (``trip free'' device 
in which the tripping mechanism cannot be overridden by the operating 
control) must be designed so that--
    (1) A manual operation is required to restore service after 
tripping; and
    (2) If an overload or circuit fault exists, the device will open the 
circuit regardless of the position of the operating control.
    (d) If the ability to reset a circuit breaker or replace a fuse is 
essential to safety in flight, that circuit breaker or fuse must be so 
located and identified that it can be readily reset or replaced in 
flight.
    (e) For fuses identified as replaceable in flight--
    (1) There must be one spare of each rating or 50 percent spare fuses 
of each rating, whichever is greater; and
    (2) The spare fuse(s) must be readily accessible to any required 
pilot.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-20, 42 FR 36969, July 18, 1977]; Amdt. 23-43, 58 FR 
18976, Apr. 9, 1993

[[Page 309]]



Sec. 23.1359  Electrical system fire protection.

    (a) Each component of the electrical system must meet the applicable 
fire protection requirements of Sec. Sec. 23.863 and 23.1182.
    (b) Electrical cables, terminals, and equipment in designated fire 
zones that are used during emergency procedures must be fire-resistant.
    (c) Insulation on electrical wire and electrical cable must be self-
extinguishing when tested at an angle of 60 degrees in accordance with 
the applicable portions of appendix F of this part, or other approved 
equivalent methods. The average burn length must not exceed 3 inches (76 
mm) and the average flame time after removal of the flame source must 
not exceed 30 seconds. Drippings from the test specimen must not 
continue to flame for more than an average of 3 seconds after falling.

[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]



Sec. 23.1361  Master switch arrangement.

    (a) There must be a master switch arrangement to allow ready 
disconnection of each electric power source from power distribution 
systems, except as provided in paragraph (b) of this section. The point 
of disconnection must be adjacent to the sources controlled by the 
switch arrangement. If separate switches are incorporated into the 
master switch arrangement, a means must be provided for the switch 
arrangement to be operated by one hand with a single movement.
    (b) Load circuits may be connected so that they remain energized 
when the master switch is open, if the circuits are isolated, or 
physically shielded, to prevent their igniting flammable fluids or 
vapors that might be liberated by the leakage or rupture of any 
flammable fluid system; and
    (1) The circuits are required for continued operation of the engine; 
or
    (2) The circuits are protected by circuit protective devices with a 
rating of five amperes or less adjacent to the electric power source.
    (3) In addition, two or more circuits installed in accordance with 
the requirements of paragraph (b)(2) of this section must not be used to 
supply a load of more than five amperes.
    (c) The master switch or its controls must be so installed that the 
switch is easily discernible and accessible to a crewmember.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-20, 42 FR 36969, July 18, 1977; Amdt. 23-43, 58 FR 
18977, Apr. 9, 1993; Amdt. 23-49, 61 FR 5169, Feb. 9, 1996]



Sec. 23.1365  Electric cables and equipment.

    (a) Each electric connecting cable must be of adequate capacity.
    (b) Any equipment that is associated with any electrical cable 
installation and that would overheat in the event of circuit overload or 
fault must be flame resistant. That equipment and the electrical cables 
must not emit dangerous quantities of toxic fumes.
    (c) Main power cables (including generator cables) in the fuselage 
must be designed to allow a reasonable degree of deformation and 
stretching without failure and must--
    (1) Be separated from flammable fluid lines; or
    (2) Be shrouded by means of electrically insulated flexible conduit, 
or equivalent, which is in addition to the normal cable insulation.
    (d) Means of identification must be provided for electrical cables, 
terminals, and connectors.
    (e) Electrical cables must be installed such that the risk of 
mechanical damage and/or damage cased by fluids vapors, or sources of 
heat, is minimized.
    (f) Where a cable cannot be protected by a circuit protection device 
or other overload protection, it must not cause a fire hazard under 
fault conditions.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 
38 FR 31824, Nov. 19, 1973; Amdt. 23-43, 58 FR 18977, Apr. 9, 1993; 
Amdt. 23-49, 61 FR 5169, Feb. 9, 1996]



Sec. 23.1367  Switches.

    Each switch must be--
    (a) Able to carry its rated current;
    (b) Constructed with enough distance or insulating material between 
current carrying parts and the housing so that vibration in flight will 
not cause shorting;
    (c) Accessible to appropriate flight crewmembers; and

[[Page 310]]

    (d) Labeled as to operation and the circuit controlled.

                                 Lights



Sec. 23.1381  Instrument lights.

    The instrument lights must--
    (a) Make each instrument and control easily readable and 
discernible;
    (b) Be installed so that their direct rays, and rays reflected from 
the windshield or other surface, are shielded from the pilot's eyes; and
    (c) Have enough distance or insulating material between current 
carrying parts and the housing so that vibration in flight will not 
cause shorting.

A cabin dome light is not an instrument light.



Sec. 23.1383  Taxi and landing lights.

    Each taxi and landing light must be designed and installed so that:
    (a) No dangerous glare is visible to the pilots.
    (b) The pilot is not seriously affected by halation.
    (c) It provides enough light for night operations.
    (d) It does not cause a fire hazard in any configuration.

[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]



Sec. 23.1385  Position light system installation.

    (a) General. Each part of each position light system must meet the 
applicable requirements of this section and each system as a whole must 
meet the requirements of Sec. Sec. 23.1387 through 23.1397.
    (b) Left and right position lights. Left and right position lights 
must consist of a red and a green light spaced laterally as far apart as 
practicable and installed on the airplane such that, with the airplane 
in the normal flying position, the red light is on the left side and the 
green light is on the right side.
    (c) Rear position light. The rear position light must be a white 
light mounted as far aft as practicable on the tail or on each wing tip.
    (d) Light covers and color filters. Each light cover or color filter 
must be at least flame resistant and may not change color or shape or 
lose any appreciable light transmission during normal use.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 
41 FR 55465, Dec. 20, 1976; Amdt. 23-43, 58 FR 18977, Apr. 9, 1993]



Sec. 23.1387  Position light system dihedral angles.

    (a) Except as provided in paragraph (e) of this section, each 
position light must, as installed, show unbroken light within the 
dihedral angles described in this section.
    (b) Dihedral angle L (left) is formed by two intersecting vertical 
planes, the first parallel to the longitudinal axis of the airplane, and 
the other at 110 degrees to the left of the first, as viewed when 
looking forward along the longitudinal axis.
    (c) Dihedral angle R (right) is formed by two intersecting vertical 
planes, the first parallel to the longitudinal axis of the airplane, and 
the other at 110 degrees to the right of the first, as viewed when 
looking forward along the longitudinal axis.
    (d) Dihedral angle A (aft) is formed by two intersecting vertical 
planes making angles of 70 degrees to the right and to the left, 
respectively, to a vertical plane passing through the longitudinal axis, 
as viewed when looking aft along the longitudinal axis.
    (e) If the rear position light, when mounted as far aft as 
practicable in accordance with Sec. 23.1385(c), cannot show unbroken 
light within dihedral angle A (as defined in paragraph (d) of this 
section), a solid angle or angles of obstructed visibility totaling not 
more than 0.04 steradians is allowable within that dihedral angle, if 
such solid angle is within a cone whose apex is at the rear position 
light and whose elements make an angle of 30[deg] with a vertical line 
passing through the rear position light.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-12, 36 FR 21278, Nov. 5, 1971; Amdt. 23-43, 58 FR 
18977, Apr. 9, 1993]



Sec. 23.1389  Position light distribution and intensities.

    (a) General. The intensities prescribed in this section must be 
provided by new equipment with each light cover and color filter in 
place. Intensities must

[[Page 311]]

be determined with the light source operating at a steady value equal to 
the average luminous output of the source at the normal operating 
voltage of the airplane. The light distribution and intensity of each 
position light must meet the requirements of paragraph (b) of this 
section.
    (b) Position lights. The light distribution and intensities of 
position lights must be expressed in terms of minimum intensities in the 
horizontal plane, minimum intensities in any vertical plane, and maximum 
intensities in overlapping beams, within dihedral angles L, R, and A, 
and must meet the following requirements:
    (1) Intensities in the horizontal plane. Each intensity in the 
horizontal plane (the plane containing the longitudinal axis of the 
airplane and perpendicular to the plane of symmetry of the airplane) 
must equal or exceed the values in Sec. 23.1391.
    (2) Intensities in any vertical plane. Each intensity in any 
vertical plane (the plane perpendicular to the horizontal plane) must 
equal or exceed the appropriate value in Sec. 23.1393, where I is the 
minimum intensity prescribed in Sec. 23.1391 for the corresponding 
angles in the horizontal plane.
    (3) Intensities in overlaps between adjacent signals. No intensity 
in any overlap between adjacent signals may exceed the values in Sec. 
23.1395, except that higher intensities in overlaps may be used with 
main beam intensities substantially greater than the minima specified in 
Sec. Sec. 23.1391 and 23.1393, if the overlap intensities in relation 
to the main beam intensities do not adversely affect signal clarity. 
When the peak intensity of the left and right position lights is more 
than 100 candles, the maximum overlap intensities between them may 
exceed the values in Sec. 23.1395 if the overlap intensity in Area A is 
not more than 10 percent of peak position light intensity and the 
overlap intensity in Area B is not more than 2.5 percent of peak 
position light intensity.
    (c) Rear position light installation. A single rear position light 
may be installed in a position displaced laterally from the plane of 
symmetry of an airplane if--
    (1) The axis of the maximum cone of illumination is parallel to the 
flight path in level flight; and
    (2) There is no obstruction aft of the light and between planes 70 
degrees to the right and left of the axis of maximum illumination.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18977, Apr. 9, 1993]



Sec. 23.1391  Minimum intensities in the horizontal plane of position lights.

    Each position light intensity must equal or exceed the applicable 
values in the following table:

------------------------------------------------------------------------
                                        Angle from right or
                                       left of longitudinal    Intensity
   Dihedral angle (light included)      axis, measured from    (candles)
                                            dead ahead
------------------------------------------------------------------------
L and R (red and green).............  0[deg] to 10[deg].....          40
                                      10[deg] to 20[deg]....          30
                                      20[deg] to 110[deg]...           5
A (rear white)......................  110[deg] to 180[deg]..          20
------------------------------------------------------------------------


[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18977, Apr. 9, 1993]



Sec. 23.1393  Minimum intensities in any vertical plane of position lights.

    Each position light intensity must equal or exceed the applicable 
values in the following table:

------------------------------------------------------------------------
                                                              Intensity,
          Angle above or below the horizontal plane                l
------------------------------------------------------------------------
0[deg]......................................................        1.00
0[deg] to 5[deg]............................................        0.90
5[deg] to 10[deg]...........................................        0.80
10[deg] to 15[deg]..........................................        0.70
15[deg] to 20[deg]..........................................        0.50
20[deg] to 30[deg]..........................................        0.30
30[deg] to 40[deg]..........................................        0.10
40[deg] to 90[deg]..........................................        0.05
------------------------------------------------------------------------


[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18977, Apr. 9, 1993]



Sec. 23.1395  Maximum intensities in overlapping beams of position lights.

    No position light intensity may exceed the applicable values in the 
following equal or exceed the applicable values in Sec. 23.1389(b)(3):

------------------------------------------------------------------------
                                                     Maximum intensity
                                                 -----------------------
                    Overlaps                        Area A      Area B
                                                   (candles)   (candles)
------------------------------------------------------------------------
Green in dihedral angle L.......................          10           1
Red in dihedral angle R.........................          10           1

[[Page 312]]

 
Green in dihedral angle A.......................           5           1
Red in dihedral angle A.........................           5           1
Rear white in dihedral angle L..................           5           1
Rear white in dihedral angle R..................           5           1
------------------------------------------------------------------------


Where--
    (a) Area A includes all directions in the adjacent dihedral angle 
that pass through the light source and intersect the common boundary 
plane at more than 10 degrees but less than 20 degrees; and
    (b) Area B includes all directions in the adjacent dihedral angle 
that pass through the light source and intersect the common boundary 
plane at more than 20 degrees.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 
58 FR 18977, Apr. 9, 1993]



Sec. 23.1397  Color specifications.

    Each position light color must have the applicable International 
Commission on Illumination chromaticity coordinates as follows:
    (a) Aviation red--

y is not greater than 0.335; and
z is not greater than 0.002.

    (b) Aviation green--

x is not greater than 0.440-0.320y;
x is not greater than y-0.170; and
y is not less than 0.390-0.170x.

    (c) Aviation white--

x is not less than 0.300 and not greater than 0.540;
y is not less than x-0.040 or y0-0.010, whichever is the 
          smaller; and
y is not greater than x+0.020 nor 0.636-0.400x;

Where y0 is the y coordinate of the Planckian radiator for 
          the value of x considered.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, amended by Amdt. 23-11, 36 
FR 12971, July 10, 1971]



Sec. 23.1399  Riding light.

    (a) Each riding (anchor) light required for a seaplane or amphibian, 
must be installed so that it can--
    (1) Show a white light for at least two miles at night under clear 
atmospheric conditions; and
    (2) Show the maximum unbroken light practicable when the airplane is 
moored or drifting on the water.
    (b) Externally hung lights may be used.



Sec. 23.1401  Anticollision light system.

    (a) General. The airplane must have an anticollision light system 
that:
    (1) Consists of one or more approved anticollision lights located so 
that their light will not impair the flight crewmembers' vision or 
detract from the conspicuity of the position lights; and
    (2) Meets the requirements of paragraphs (b) through (f) of this 
section.
    (b) Field of coverage. The system must consist of enough lights to 
illuminate the vital areas around the airplane, considering the physical 
configuration and flight characteristics of the airplane. The field of 
coverage must extend in each direction within at least 75 degrees above 
and 75 degrees below the horizontal plane of the airplane, except that 
there may be solid angles of obstructed visibility totaling not more 
than 0.5 steradians.
    (c) Flashing characteristics. The arrangement of the system, that 
is, the number of light sources, beam width, speed of rotation, and 
other characteristics, must give an effective flash frequency of not 
less than 40, nor more than 100, cycles per minute. The effective flash 
frequency is the frequency at which the airplane's complete 
anticollision light system is observed from a distance, and applies to 
each sector of light including any overlaps that exist when the system 
consists of more than one light source. In overlaps, flash frequencies 
may exceed 100, but not 180, cycles per minute.
    (d) Color. Each anticollision light must be either aviation red or 
aviation white and must meet the applicable requirements of Sec. 
23.1397.
    (e) Light intensity. The minimum light intensities in any vertical 
plane, measured with the red filter (if used) and expressed in terms of 
``effective'' intensities, must meet the requirements of paragraph (f) 
of this section. The following relation must be assumed:

[[Page 313]]

[GRAPHIC] [TIFF OMITTED] TC28SE91.018

where:

Ie=effective intensity (candles).
I(t)=instantaneous intensity as a function of time.
t2-t1=flash time interval (seconds).


Normally, the maximum value of effective intensity is obtained when 
t2 and t1 are chosen so that the effective 
intensity is equal to the instantaneous intensity at t2 and 
t1.
    (f) Minimum effective intensities for anticollision lights. Each 
anticollision light effective intensity must equal or exceed the 
applicable values in the following table.

------------------------------------------------------------------------
                                                              Effective
         Angle above or below the horizontal plane            intensity
                                                              (candles)
------------------------------------------------------------------------
0[deg] to 5[deg]...........................................          400
5[deg] to 10[deg]..........................................          240
10[deg] to 20[deg].........................................           80
20[deg] to 30[deg].........................................           40
30[deg] to 75[deg].........................................           20
------------------------------------------------------------------------


[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-11, 
36 FR 12972, July 10, 1971; Amdt. 23-20, 42 FR 36969, July 18, 1977; 
Amdt. 23-49, 61 FR 5169, Feb. 9, 1996]

                            Safety Equipment



Sec. 23.1411  General.

    (a) Required safety equipment to be used by the flight crew in an 
emergency, such as automatic liferaft releases, must be readily 
accessible.
    (b) Stowage provisions for required safety equipment must be 
furnished and must--
    (1) Be arranged so that the equipment is directly accessible and its 
location is obvious; and
    (2) Protect the safety equipment from damage caused by being 
subjected to the inertia loads resulting from the ultimate static load 
factors specified in Sec. 23.561(b)(3) of this part.

[Amdt. 23-17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23-36, 53 
FR 30815, Aug. 15, 1988]



Sec. 23.1415  Ditching equipment.

    (a) Emergency flotation and signaling equipment required by any 
operating rule in this chapter must be installed so that it is readily 
available to the crew and passengers.
    (b) Each raft and each life preserver must be approved.
    (c) Each raft released automatically or by the pilot must be 
attached to the airplane by a line to keep it alongside the airplane. 
This line must be weak enough to break before submerging the empty raft 
to which it is attached.
    (d) Each signaling device required by any operating rule in this 
chapter, must be accessible, function satisfactorily, and must be free 
of any hazard in its operation.



Sec. 23.1416  Pneumatic de-icer boot system.

    If certification with ice protection provisions is desired and a 
pneumatic de-icer boot system is installed--
    (a) The system must meet the requirements specified in Sec. 
23.1419.
    (b) The system and its components must be designed to perform their 
intended function under any normal system operating temperature or 
pressure, and
    (c) Means to indicate to the flight crew that the pneumatic de-icer 
boot system is receiving adequate pressure and is functioning normally 
must be provided.

[Amdt. 23-23, 43 FR 50593, Oct. 30, 1978]



Sec. 23.1419  Ice protection.

    If certification with ice protection provisions is desired, 
compliance with the requirements of this section and other applicable 
sections of this part must be shown:
    (a) An analysis must be performed to establish, on the basis of the 
airplane's operational needs, the adequacy of the ice protection system 
for the various components of the airplane. In addition, tests of the 
ice protection system must be conducted to demonstrate that the airplane 
is capable of operating safely in continuous maximum and intermittent 
maximum icing conditions, as described in appendix C of part 25 of this 
chapter. As used in this section, ``Capable of operating safely,'' means 
that airplane performance, controllability, maneuverability, and 
stability must not be less than that required in part 23, subpart B.

[[Page 314]]

    (b) Except as provided by paragraph (c) of this section, in addition 
to the analysis and physical evaluation prescribed in paragraph (a) of 
this section, the effectiveness of the ice protection system and its 
components must be shown by flight tests of the airplane or its 
components in measured natural atmospheric icing conditions and by one 
or more of the following tests, as found necessary to determine the 
adequacy of the ice protection system--
    (1) Laboratory dry air or simulated icing tests, or a combination of 
both, of the components or models of the components.
    (2) Flight dry air tests of the ice protection system as a whole, or 
its individual components.
    (3) Flight test of the airplane or its components in measured 
simulated icing conditions.
    (c) If certification with ice protection has been accomplished on 
prior type certificated airplanes whose designs include components that 
are thermodynamically and aerodynamically equivalent to those used on a 
new airplane design, certification of these equivalent components may be 
accomplished by reference to previously accomplished tests, required in 
Sec. 23.1419 (a) and (b), provided that the applicant accounts for any 
differences in installation of these components.
    (d) A means must be identified or provided for determining the 
formation of ice on the critical parts of the airplane. Adequate 
lighting must be provided for the use of this means during night 
operation. Also, when monitoring of the external surfaces of the 
airplane by the flight crew is required for operation of the ice 
protection equipment, external lighting must be provided that is 
adequate to enable the monitoring to be done at night. Any illumination 
that is used must be of a type that will not cause glare or reflection 
that would handicap crewmembers in the performance of their duties. The 
Airplane Flight Manual or other approved manual material must describe 
the means of determining ice formation and must contain information for 
the safe operation of the airplane in icing conditions.

[Doc. No. 26344, 58 FR 18977, Apr. 9, 1993]

                         Miscellaneous Equipment



Sec. 23.1431  Electronic equipment.

    (a) In showing compliance with Sec. 23.1309(a), (b), and (c) with 
respect to radio and electronic equipment and their installations, 
critical environmental conditions must be considered.
    (b) Radio and electronic equipment, controls, and wiring must be 
installed so that operation of any unit or system of units will not 
adversely affect the simultaneous operation of any other radio or 
electronic unit, or system of units, required by this chapter.
    (c) For those airplanes required to have more than one flightcrew 
member, or whose operation will require more than one flightcrew member, 
the cockpit must be evaluated to determine if the flightcrew members, 
when seated at their duty station, can converse without difficulty under 
the actual cockpit noise conditions when the airplane is being operated. 
If the airplane design includes provision for the use of communication 
headsets, the evaluation must also consider conditions where headsets 
are being used. If the evaluation shows conditions under which it will 
be difficult to converse, an intercommunication system must be provided.
    (d) If installed communication equipment includes transmitter ``off-
on'' switching, that switching means must be designed to return from the 
``transmit'' to the ``off'' position when it is released and ensure that 
the transmitter will return to the off (non transmitting) state.
    (e) If provisions for the use of communication headsets are 
provided, it must be demonstrated that the flightcrew members will 
receive all aural warnings under the actual cockpit noise conditions 
when the airplane is being operated when any headset is being used.

[Doc. No. 26344, 58 FR 18977, Apr. 9, 1993, as amended by Amdt. 23-49, 
61 FR 5169, Feb. 9, 1996; Amdt. 23-62, 76 FR 75761, Dec. 2, 2011]



Sec. 23.1435  Hydraulic systems.

    (a) Design. Each hydraulic system must be designed as follows:

[[Page 315]]

    (1) Each hydraulic system and its elements must withstand, without 
yielding, the structural loads expected in addition to hydraulic loads.
    (2) A means to indicate the pressure in each hydraulic system which 
supplies two or more primary functions must be provided to the flight 
crew.
    (3) There must be means to ensure that the pressure, including 
transient (surge) pressure, in any part of the system will not exceed 
the safe limit above design operating pressure and to prevent excessive 
pressure resulting from fluid volumetric changes in all lines which are 
likely to remain closed long enough for such changes to occur.
    (4) The minimum design burst pressure must be 2.5 times the 
operating pressure.
    (b) Tests. Each system must be substantiated by proof pressure 
tests. When proof tested, no part of any system may fail, malfunction, 
or experience a permanent set. The proof load of each system must be at 
least 1.5 times the maximum operating pressure of that system.
    (c) Accumulators. A hydraulic accumulator or reservoir may be 
installed on the engine side of any firewall if--
    (1) It is an integral part of an engine or propeller system, or
    (2) The reservoir is nonpressurized and the total capacity of all 
such nonpressurized reservoirs is one quart or less.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13096, Aug. 13, 1969; Amdt. 23-14, 38 FR 31824, Nov. 19, 1973; Amdt. 
23-43, 58 FR 18977, Apr. 9, 1993; Amdt. 23-49, 61 FR 5170, Feb. 9, 1996]



Sec. 23.1437  Accessories for multiengine airplanes.

    For multiengine airplanes, engine-driven accessories essential to 
safe operation must be distributed among two or more engines so that the 
failure of any one engine will not impair safe operation through the 
malfunctioning of these accessories.



Sec. 23.1438  Pressurization and pneumatic systems.

    (a) Pressurization system elements must be burst pressure tested to 
2.0 times, and proof pressure tested to 1.5 times, the maximum normal 
operating pressure.
    (b) Pneumatic system elements must be burst pressure tested to 3.0 
times, and proof pressure tested to 1.5 times, the maximum normal 
operating pressure.
    (c) An analysis, or a combination of analysis and test, may be 
substituted for any test required by paragraph (a) or (b) of this 
section if the Administrator finds it equivalent to the required test.

[Amdt. 23-20, 42 FR 36969, July 18, 1977]



Sec. 23.1441  Oxygen equipment and supply.

    (a) If certification with supplemental oxygen equipment is 
requested, or the airplane is approved for operations at or above 
altitudes where oxygen is required to be used by the operating rules, 
oxygen equipment must be provided that meets the requirements of this 
section and Sec. Sec. 23.1443 through 23.1449. Portable oxygen 
equipment may be used to meet the requirements of this part if the 
portable equipment is shown to comply with the applicable requirements, 
is identified in the airplane type design, and its stowage provisions 
are found to be in compliance with the requirements of Sec. 23.561.
    (b) The oxygen system must be free from hazards in itself, in its 
method of operation, and its effect upon other components.
    (c) There must be a means to allow the crew to readily determine, 
during the flight, the quantity of oxygen available in each source of 
supply.
    (d) Each required flight crewmember must be provided with--
    (1) Demand oxygen equipment if the airplane is to be certificated 
for operation above 25,000 feet.
    (2) Pressure demand oxygen equipment if the airplane is to be 
certificated for operation above 40,000 feet.
    (e) There must be a means, readily available to the crew in flight, 
to turn on and to shut off the oxygen supply at the high pressure 
source. This shutoff requirement does not apply to chemical oxygen 
generators.

[Amdt. 23-9, 35 FR 6386, Apr. 21, 1970, as amended by Amdt. 23-43, 58 FR 
18978, Apr. 9, 1993]

[[Page 316]]



Sec. 23.1443  Minimum mass flow of supplemental oxygen.

    (a) If the airplane is to be certified above 41,000 feet, a 
continuous flow oxygen system must be provided for each passenger.
    (b) If continuous flow oxygen equipment is installed, an applicant 
must show compliance with the requirements of either paragraphs (b)(1) 
and (b)(2) or paragraph (b)(3) of this section:
    (1) For each passenger, the minimum mass flow of supplemental oxygen 
required at various cabin pressure altitudes may not be less than the 
flow required to maintain, during inspiration and while using the oxygen 
equipment (including masks) provided, the following mean tracheal oxygen 
partial pressures:
    (i) At cabin pressure altitudes above 10,000 feet up to and 
including 18,500 feet, a mean tracheal oxygen partial pressure of 100mm 
Hg when breathing 15 liters per minute, Body Temperature, Pressure, 
Saturated (BTPS) and with a tidal volume of 700cc with a constant time 
interval between respirations.
    (ii) At cabin pressure altitudes above 18,500 feet up to and 
including 40,000 feet, a mean tracheal oxygen partial pressure of 83.8mm 
Hg when breathing 30 liters per minute, BTPS, and with a tidal volume of 
1,100cc with a constant time interval between respirations.
    (2) For each flight crewmember, the minimum mass flow may not be 
less than the flow required to maintain, during inspiration, a mean 
tracheal oxygen partial pressure of 149mm Hg when breathing 15 liters 
per minute, BTPS, and with a maximum tidal volume of 700cc with a 
constant time interval between respirations.
    (3) The minimum mass flow of supplemental oxygen supplied for each 
user must be at a rate not less than that shown in the following figure 
for each altitude up to and including the maximum operating altitude of 
the airplane.
[GRAPHIC] [TIFF OMITTED] TR02DE11.086


[[Page 317]]


    (c) If demand equipment is installed for use by flight crewmembers, 
the minimum mass flow of supplemental oxygen required for each flight 
crewmember may not be less than the flow required to maintain, during 
inspiration, a mean tracheal oxygen partial pressure of 122mm Hg up to 
and including a cabin pressure altitude of 35,000 feet, and 95 percent 
oxygen between cabin pressure altitudes of 35,000 and 40,000 feet, when 
breathing 20 liters per minutes BTPS. In addition, there must be means 
to allow the flight crew to use undiluted oxygen at their discretion.
    (d) If first-aid oxygen equipment is installed, the minimum mass 
flow of oxygen to each user may not be less than 4 liters per minute, 
STPD. However, there may be a means to decrease this flow to not less 
than 2 liters per minute, STPD, at any cabin altitude. The quantity of 
oxygen required is based upon an average flow rate of 3 liters per 
minute per person for whom first-aid oxygen is required.
    (e) As used in this section:
    (1) BTPS means Body Temperature, and Pressure, Saturated (which is 
37 [deg]C, and the ambient pressure to which the body is exposed, minus 
47mm Hg, which is the tracheal pressure displaced by water vapor 
pressure when the breathed air becomes saturated with water vapor at 37 
[deg]C).
    (2) STPD means Standard, Temperature, and Pressure, Dry (which is 0 
[deg]C at 760mm Hg with no water vapor).

[Doc. No. FAA-2009-0738, 76 FR 75761, Dec. 2, 2011]



Sec. 23.1445  Oxygen distribution system.

    (a) Except for flexible lines from oxygen outlets to the dispensing 
units, or where shown to be otherwise suitable to the installation, 
nonmetallic tubing must not be used for any oxygen line that is normally 
pressurized during flight.
    (b) Nonmetallic oxygen distribution lines must not be routed where 
they may be subjected to elevated temperatures, electrical arcing, and 
released flammable fluids that might result from any probable failure.
    (c) If the flight crew and passengers share a common source of 
oxygen, a means to separately reserve the minimum supply required by the 
flight crew must be provided.

[Doc. No. 26344, 58 FR 18978, Apr. 9, 1993, as amended by Amdt. 23-62, 
76 FR 75762, Dec. 2, 2011]



Sec. 23.1447  Equipment standards for oxygen dispensing units.

    If oxygen dispensing units are installed, the following apply:
    (a) There must be an individual dispensing unit for each occupant 
for whom supplemental oxygen is to be supplied. Each dispensing unit 
must:
    (1) Provide for effective utilization of the oxygen being delivered 
to the unit.
    (2) Be capable of being readily placed into position on the face of 
the user.
    (3) Be equipped with a suitable means to retain the unit in position 
on the face.
    (4) If radio equipment is installed, the flightcrew oxygen 
dispensing units must be designed to allow the use of that equipment and 
to allow communication with any other required crew member while at 
their assigned duty station.
    (b) If certification for operation up to and including 18,000 feet 
(MSL) is requested, each oxygen dispensing unit must:
    (1) Cover the nose and mouth of the user; or
    (2) Be a nasal cannula, in which case one oxygen dispensing unit 
covering both the nose and mouth of the user must be available. In 
addition, each nasal cannula or its connecting tubing must have 
permanently affixed--
    (i) A visible warning against smoking while in use;
    (ii) An illustration of the correct method of donning; and
    (iii) A visible warning against use with nasal obstructions or head 
colds with resultant nasal congestion.
    (c) If certification for operation above 18,000 feet (MSL) is 
requested, each oxygen dispensing unit must cover the nose and mouth of 
the user.
    (d) For a pressurized airplane designed to operate at flight 
altitudes above 25,000 feet (MSL), the dispensing units must meet the 
following:

[[Page 318]]

    (1) The dispensing units for passengers must be connected to an 
oxygen supply terminal and be immediately available to each occupant 
wherever seated.
    (2) The dispensing units for crewmembers must be automatically 
presented to each crewmember before the cabin pressure altitude exceeds 
15,000 feet, or the units must be of the quick-donning type, connected 
to an oxygen supply terminal that is immediately available to 
crewmembers at their station.
    (e) If certification for operation above 30,000 feet is requested, 
the dispensing units for passengers must be automatically presented to 
each occupant before the cabin pressure altitude exceeds 15,000 feet.
    (f) If an automatic dispensing unit (hose and mask, or other unit) 
system is installed, the crew must be provided with a manual means to 
make the dispensing units immediately available in the event of failure 
of the automatic system.
    (g) If the airplane is to be certified for operation above 41,000 
feet, a quick-donning oxygen mask system, with a pressure demand, mask 
mounted regulator must be provided for the flight crew. This dispensing 
unit must be immediately available to the flight crew when seated at 
their station and installed so that it:
    (1) Can be placed on the face from its ready position, properly 
secured, sealed, and supplying oxygen upon demand, with one hand, within 
five seconds and without disturbing eyeglasses or causing delay in 
proceeding with emergency duties; and
    (2) Allows, while in place, the performance of normal communication 
functions.

[Amdt. 23-9, 35 FR 6387, Apr. 21, 1970, as amended by Amdt. 23-20, 42 FR 
36969, July 18, 1977; Amdt. 23-30, 49 FR 7340, Feb. 28, 1984; Amdt. 23-
43, 58 FR 18978, Apr. 9, 1993; Amdt. 23-49, 61 FR 5170, Feb. 9, 1996; 
Amdt. 23-62, 76 FR 75762, Dec. 2, 2011]



Sec. 23.1449  Means for determining use of oxygen.

    There must be a means to allow the crew to determine whether oxygen 
is being delivered to the dispensing equipment.

[Amdt. 23-9, 35 FR 6387, Apr. 21, 1970]



Sec. 23.1450  Chemical oxygen generators.

    (a) For the purpose of this section, a chemical oxygen generator is 
defined as a device which produces oxygen by chemical reaction.
    (b) Each chemical oxygen generator must be designed and installed in 
accordance with the following requirements:
    (1) Surface temperature developed by the generator during operation 
may not create a hazard to the airplane or to its occupants.
    (2) Means must be provided to relieve any internal pressure that may 
be hazardous.
    (c) In addition to meeting the requirements in paragraph (b) of this 
section, each portable chemical oxygen generator that is capable of 
sustained operation by successive replacement of a generator element 
must be placarded to show--
    (1) The rate of oxygen flow, in liters per minute;
    (2) The duration of oxygen flow, in minutes, for the replaceable 
generator element; and
    (3) A warning that the replaceable generator element may be hot, 
unless the element construction is such that the surface temperature 
cannot exceed 100 [deg]F.

[Amdt. 23-20, 42 FR 36969, July 18, 1977]



Sec. 23.1451  Fire protection for oxygen equipment.

    Oxygen equipment and lines must:
    (a) Not be installed in any designed fire zones.
    (b) Be protected from heat that may be generated in, or escape from, 
any designated fire zone.
    (c) Be installed so that escaping oxygen cannot come in contact with 
and cause ignition of grease, fluid, or vapor accumulations that are 
present in normal operation or that may result from the failure or 
malfunction of any other system.

[Doc. No. 27806, 61 FR 5170, Feb. 9, 1996]

[[Page 319]]



Sec. 23.1453  Protection of oxygen equipment from rupture.

    (a) Each element of the oxygen system must have sufficient strength 
to withstand the maximum pressure and temperature, in combination with 
any externally applied loads arising from consideration of limit 
structural loads, that may be acting on that part of the system.
    (b) Oxygen pressure sources and the lines between the source and the 
shutoff means must be:
    (1) Protected from unsafe temperatures; and
    (2) Located where the probability and hazard of rupture in a crash 
landing are minimized.

[Doc. No. 27806, 61 FR 5170, Feb. 9, 1996]



Sec. 23.1457  Cockpit voice recorders.

    (a) Each cockpit voice recorder required by the operating rules of 
this chapter must be approved and must be installed so that it will 
record the following:
    (1) Voice communications transmitted from or received in the 
airplane by radio.
    (2) Voice communications of flight crewmembers on the flight deck.
    (3) Voice communications of flight crewmembers on the flight deck, 
using the airplane's interphone system.
    (4) Voice or audio signals identifying navigation or approach aids 
introduced into a headset or speaker.
    (5) Voice communications of flight crewmembers using the passenger 
loudspeaker system, if there is such a system and if the fourth channel 
is available in accordance with the requirements of paragraph (c)(4)(ii) 
of this section.
    (6) If datalink communication equipment is installed, all datalink 
communications, using an approved data message set. Datalink messages 
must be recorded as the output signal from the communications unit that 
translates the signal into usable data.
    (b) The recording requirements of paragraph (a)(2) of this section 
must be met by installing a cockpit-mounted area microphone, located in 
the best position for recording voice communications originating at the 
first and second pilot stations and voice communications of other 
crewmembers on the flight deck when directed to those stations. The 
microphone must be so located and, if necessary, the preamplifiers and 
filters of the recorder must be so adjusted or supplemented, so that the 
intelligibility of the recorded communications is as high as practicable 
when recorded under flight cockpit noise conditions and played back. 
Repeated aural or visual playback of the record may be used in 
evaluating intelligibility.
    (c) Each cockpit voice recorder must be installed so that the part 
of the communication or audio signals specified in paragraph (a) of this 
section obtained from each of the following sources is recorded on a 
separate channel:
    (1) For the first channel, from each boom, mask, or handheld 
microphone, headset, or speaker used at the first pilot station.
    (2) For the second channel from each boom, mask, or handheld 
microphone, headset, or speaker used at the second pilot station.
    (3) For the third channel--from the cockpit-mounted area microphone.
    (4) For the fourth channel from:
    (i) Each boom, mask, or handheld microphone, headset, or speaker 
used at the station for the third and fourth crewmembers.
    (ii) If the stations specified in paragraph (c)(4)(i) of this 
section are not required or if the signal at such a station is picked up 
by another channel, each microphone on the flight deck that is used with 
the passenger loudspeaker system, if its signals are not picked up by 
another channel.
    (5) And that as far as is practicable all sounds received by the 
microphone listed in paragraphs (c)(1), (2), and (4) of this section 
must be recorded without interruption irrespective of the position of 
the interphone-transmitter key switch. The design shall ensure that 
sidetone for the flight crew is produced only when the interphone, 
public address system, or radio transmitters are in use.
    (d) Each cockpit voice recorder must be installed so that:
    (1)(i) It receives its electrical power from the bus that provides 
the maximum reliability for operation of the

[[Page 320]]

cockpit voice recorder without jeopardizing service to essential or 
emergency loads.
    (ii) It remains powered for as long as possible without jeopardizing 
emergency operation of the airplane.
    (2) There is an automatic means to simultaneously stop the recorder 
and prevent each erasure feature from functioning, within 10 minutes 
after crash impact; and
    (3) There is an aural or visual means for preflight checking of the 
recorder for proper operation;
    (4) Any single electrical failure external to the recorder does not 
disable both the cockpit voice recorder and the flight data recorder;
    (5) It has an independent power source--
    (i) That provides 10 1 minutes of electrical 
power to operate both the cockpit voice recorder and cockpit-mounted 
area microphone;
    (ii) That is located as close as practicable to the cockpit voice 
recorder; and
    (iii) To which the cockpit voice recorder and cockpit-mounted area 
microphone are switched automatically in the event that all other power 
to the cockpit voice recorder is interrupted either by normal shutdown 
or by any other loss of power to the electrical power bus; and
    (6) It is in a separate container from the flight data recorder when 
both are required. If used to comply with only the cockpit voice 
recorder requirements, a combination unit may be installed.
    (e) The recorder container must be located and mounted to minimize 
the probability of rupture of the container as a result of crash impact 
and consequent heat damage to the recorder from fire.
    (1) Except as provided in paragraph (e)(2) of this section, the 
recorder container must be located as far aft as practicable, but need 
not be outside of the pressurized compartment, and may not be located 
where aft-mounted engines may crush the container during impact.
    (2) If two separate combination digital flight data recorder and 
cockpit voice recorder units are installed instead of one cockpit voice 
recorder and one digital flight data recorder, the combination unit that 
is installed to comply with the cockpit voice recorder requirements may 
be located near the cockpit.
    (f) If the cockpit voice recorder has a bulk erasure device, the 
installation must be designed to minimize the probability of inadvertent 
operation and actuation of the device during crash impact.
    (g) Each recorder container must:
    (1) Be either bright orange or bright yellow;
    (2) Have reflective tape affixed to its external surface to 
facilitate its location under water; and
    (3) Have an underwater locating device, when required by the 
operating rules of this chapter, on or adjacent to the container which 
is secured in such manner that they are not likely to be separated 
during crash impact.

[Amdt. 23-35, 53 FR 26142, July 11, 1988, as amended by Amdt. 23-58, 73 
FR 12562, Mar. 7, 2008; 74 FR 32799, July 9, 2009]



Sec. 23.1459  Flight data recorders.

    (a) Each flight recorder required by the operating rules of this 
chapter must be installed so that:
    (1) It is supplied with airspeed, altitude, and directional data 
obtained from sources that meet the accuracy requirements of Sec. Sec. 
23.1323, 23.1325, and 23.1327, as appropriate;
    (2) The vertical acceleration sensor is rigidly attached, and 
located longitudinally either within the approved center of gravity 
limits of the airplane, or at a distance forward or aft of these limits 
that does not exceed 25 percent of the airplane's mean aerodynamic 
chord;
    (3)(i) It receives its electrical power from the bus that provides 
the maximum reliability for operation of the flight data recorder 
without jeopardizing service to essential or emergency loads.
    (ii) It remains powered for as long as possible without jeopardizing 
emergency operation of the airplane.
    (4) There is an aural or visual means for preflight checking of the 
recorder for proper recording of data in the storage medium;

[[Page 321]]

    (5) Except for recorders powered solely by the engine-driven 
electrical generator system, there is an automatic means to 
simultaneously stop a recorder that has a data erasure feature and 
prevent each erasure feature from functioning, within 10 minutes after 
crash impact;
    (6) Any single electrical failure external to the recorder does not 
disable both the cockpit voice recorder and the flight data recorder; 
and
    (7) It is in a separate container from the cockpit voice recorder 
when both are required. If used to comply with only the flight data 
recorder requirements, a combination unit may be installed. If a 
combination unit is installed as a cockpit voice recorder to comply with 
Sec. 23.1457(e)(2), a combination unit must be used to comply with this 
flight data recorder requirement.
    (b) Each nonejectable record container must be located and mounted 
so as to minimize the probability of container rupture resulting from 
crash impact and subsequent damage to the record from fire. In meeting 
this requirement the record container must be located as far aft as 
practicable, but need not be aft of the pressurized compartment, and may 
not be where aft-mounted engines may crush the container upon impact.
    (c) A correlation must be established between the flight recorder 
readings of airspeed, altitude, and heading and the corresponding 
readings (taking into account correction factors) of the first pilot's 
instruments. The correlation must cover the airspeed range over which 
the airplane is to be operated, the range of altitude to which the 
airplane is limited, and 360 degrees of heading. Correlation may be 
established on the ground as appropriate.
    (d) Each recorder container must:
    (1) Be either bright orange or bright yellow;
    (2) Have reflective tape affixed to its external surface to 
facilitate its location under water; and
    (3) Have an underwater locating device, when required by the 
operating rules of this chapter, on or adjacent to the container which 
is secured in such a manner that they are not likely to be separated 
during crash impact.
    (e) Any novel or unique design or operational characteristics of the 
aircraft shall be evaluated to determine if any dedicated parameters 
must be recorded on flight recorders in addition to or in place of 
existing requirements.

[Amdt. 23-35, 53 FR 26143, July 11, 1988, as amended by Amdt. 23-58, 73 
FR 12562, Mar. 7, 2008; 74 FR 32800, July 9, 2009]



Sec. 23.1461  Equipment containing high energy rotors.

    (a) Equipment, such as Auxiliary Power Units (APU) and constant 
speed drive units, containing high energy rotors must meet paragraphs 
(b), (c), or (d) of this section.
    (b) High energy rotors contained in equipment must be able to 
withstand damage caused by malfunctions, vibration, abnormal speeds, and 
abnormal temperatures. In addition--
    (1) Auxiliary rotor cases must be able to contain damage caused by 
the failure of high energy rotor blades; and
    (2) Equipment control devices, systems, and instrumentation must 
reasonably ensure that no operating limitations affecting the integrity 
of high energy rotors will be exceeded in service.
    (c) It must be shown by test that equipment containing high energy 
rotors can contain any failure of a high energy rotor that occurs at the 
highest speed obtainable with the normal speed control devices 
inoperative.
    (d) Equipment containing high energy rotors must be located where 
rotor failure will neither endanger the occupants nor adversely affect 
continued safe flight.

[Amdt. 23-20, 42 FR 36969, July 18, 1977, as amended by Amdt. 23-49, 61 
FR 5170, Feb. 9, 1996]



             Subpart G_Operating Limitations and Information



Sec. 23.1501  General.

    (a) Each operating limitation specified in Sec. Sec. 23.1505 
through 23.1527 and other limitations and information necessary for safe 
operation must be established.
    (b) The operating limitations and other information necessary for 
safe operation must be made available to

[[Page 322]]

the crewmembers as prescribed in Sec. Sec. 23.1541 through 23.1589.

[Amdt. 23-21, 43 FR 2319, Jan. 16, 1978]



Sec. 23.1505  Airspeed limitations.

    (a) The never-exceed speed VNE must be established so that it is--
    (1) Not less than 0.9 times the minimum value of VD allowed under 
Sec. 23.335; and
    (2) Not more than the lesser of--
    (i) 0.9 VD established under Sec. 23.335; or
    (ii) 0.9 times the maximum speed shown under Sec. 23.251.
    (b) The maximum structural cruising speed VNO must be established so 
that it is--
    (1) Not less than the minimum value of VC allowed under Sec. 
23.335; and
    (2) Not more than the lesser of--
    (i) VC established under Sec. 23.335; or
    (ii) 0.89 VNE established under paragraph (a) of this section.
    (c)(1) Paragraphs (a) and (b) of this section do not apply to 
turbine airplanes or to airplanes for which a design diving speed 
VD/MD is established under Sec. 23.335(b)(4). For 
those airplanes, a maximum operating limit speed (VMO/
MMO airspeed or Mach number, whichever is critical at a 
particular altitude) must be established as a speed that may not be 
deliberately exceeded in any regime of flight (climb, cruise, or 
descent) unless a higher speed is authorized for flight test or pilot 
training operations.
    (2) VMO/MMO must be established so that it is 
not greater than the design cruising speed VC/MC 
and so that it is sufficiently below VD/MD, or 
VDF/MDF for jets, and the maximum speed shown 
under Sec. 23.251 to make it highly improbable that the latter speeds 
will be inadvertently exceeded in operations.
    (3) The speed margin between VMO/MMO and 
VD/MD, or VDF/MDF for jets, 
may not be less than that determined under Sec. 23.335(b), or the speed 
margin found necessary in the flight tests conducted under Sec. 23.253.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13096, Aug. 13, 1969; Amdt. 23-62, 76 FR 75762, Dec. 2, 2011]



Sec. 23.1507  Operating maneuvering speed.

    The maximum operating maneuvering speed, VO, must be 
established as an operating limitation. VO is a selected 
speed that is not greater than VS[radic]n established in 
Sec. 23.335(c).

[Doc. No. 26269, 58 FR 42165, Aug. 6, 1993]



Sec. 23.1511  Flap extended speed.

    (a) The flap extended speed VFE must be established so that it is--
    (1) Not less than the minimum value of VF allowed in 
Sec. 23.345(b); and
    (2) Not more than VF established under Sec. 23.345(a), 
(c), and (d).
    (b) Additional combinations of flap setting, airspeed, and engine 
power may be established if the structure has been proven for the 
corresponding design conditions.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-50, 61 FR 5192, Feb. 9, 1996]



Sec. 23.1513  Minimum control speed.

    The minimum control speed VMC, determined under Sec. 23.149, must 
be established as an operating limitation.



Sec. 23.1519  Weight and center of gravity.

    The weight and center of gravity limitations determined under Sec. 
23.23 must be established as operating limitations.



Sec. 23.1521  Powerplant limitations.

    (a) General. The powerplant limitations prescribed in this section 
must be established so that they do not exceed the corresponding limits 
for which the engines or propellers are type certificated. In addition, 
other powerplant limitations used in determining compliance with this 
part must be established.
    (b) Takeoff operation. The powerplant takeoff operation must be 
limited by--
    (1) The maximum rotational speed (rpm);
    (2) The maximum allowable manifold pressure (for reciprocating 
engines);
    (3) The maximum allowable gas temperature (for turbine engines);
    (4) The time limit for the use of the power or thrust corresponding 
to the limitations established in paragraphs (b)(1) through (3) of this 
section; and
    (5) The maximum allowable cylinder head (as applicable), liquid 
coolant and oil temperatures.
    (c) Continuous operation. The continuous operation must be limited 
by--

[[Page 323]]

    (1) The maximum rotational speed;
    (2) The maximum allowable manifold pressure (for reciprocating 
engines);
    (3) The maximum allowable gas temperature (for turbine engines); and
    (4) The maximum allowable cylinder head, oil, and liquid coolant 
temperatures.
    (d) Fuel grade or designation. The minimum fuel grade (for 
reciprocating engines), or fuel designation (for turbine engines), must 
be established so that it is not less than that required for the 
operation of the engines within the limitations in paragraphs (b) and 
(c) of this section.
    (e) Ambient temperature. For all airplanes except reciprocating 
engine-powered airplanes of 6,000 pounds or less maximum weight, ambient 
temperature limitations (including limitations for winterization 
installations if applicable) must be established as the maximum ambient 
atmospheric temperature at which compliance with the cooling provisions 
of Sec. Sec. 23.1041 through 23.1047 is shown.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978; Amdt. 23-45, 58 FR 
42165, Aug. 6, 1993; Amdt. 23-50, 61 FR 5192, Feb. 9, 1996]



Sec. 23.1522  Auxiliary power unit limitations.

    If an auxiliary power unit is installed, the limitations established 
for the auxiliary power must be specified in the operating limitations 
for the airplane.

[Doc. No. 26269, 58 FR 42166, Aug. 6, 1993]



Sec. 23.1523  Minimum flight crew.

    The minimum flight crew must be established so that it is sufficient 
for safe operation considering--
    (a) The workload on individual crewmembers and, in addition for 
commuter category airplanes, each crewmember workload determination must 
consider the following:
    (1) Flight path control,
    (2) Collision avoidance,
    (3) Navigation,
    (4) Communications,
    (5) Operation and monitoring of all essential airplane systems,
    (6) Command decisions, and
    (7) The accessibility and ease of operation of necessary controls by 
the appropriate crewmember during all normal and emergency operations 
when at the crewmember flight station;
    (b) The accessibility and ease of operation of necessary controls by 
the appropriate crewmember; and
    (c) The kinds of operation authorized under Sec. 23.1525.

[Amdt. 23-21, 43 FR 2319, Jan. 16, 1978, as amended by Amdt. 23-34, 52 
FR 1834, Jan. 15, 1987]



Sec. 23.1524  Maximum passenger seating configuration.

    The maximum passenger seating configuration must be established.

[Amdt. 23-10, 36 FR 2864, Feb. 11, 1971]



Sec. 23.1525  Kinds of operation.

    The kinds of operation authorized (e.g. VFR, IFR, day or night) and 
the meteorological conditions (e.g. icing) to which the operation of the 
airplane is limited or from which it is prohibited, must be established 
appropriate to the installed equipment.

[Doc. No. 26269, 58 FR 42166, Aug. 6, 1993]



Sec. 23.1527  Maximum operating altitude.

    (a) The maximum altitude up to which operation is allowed, as 
limited by flight, structural, powerplant, functional or equipment 
characteristics, must be established.
    (b) A maximum operating altitude limitation of not more than 25,000 
feet must be established for pressurized airplanes unless compliance 
with Sec. 23.775(e) is shown.

[Doc. No. 26269, 58 FR 42166, Aug. 6, 1993]



Sec. 23.1529  Instructions for Continued Airworthiness.

    The applicant must prepare Instructions for Continued Airworthiness 
in accordance with appendix G to this part that are acceptable to the 
Administrator. The instructions may be incomplete at type certification 
if a program exists to ensure their completion prior to delivery of the 
first airplane or issuance of a standard certificate of airworthiness, 
whichever occurs later.

[Amdt. 23-26, 45 FR 60171, Sept. 11, 1980]

[[Page 324]]

                          Markings And Placards



Sec. 23.1541  General.

    (a) The airplane must contain--
    (1) The markings and placards specified in Sec. Sec. 23.1545 
through 23.1567; and
    (2) Any additional information, instrument markings, and placards 
required for the safe operation if it has unusual design, operating, or 
handling characteristics.
    (b) Each marking and placard prescribed in paragraph (a) of this 
section--
    (1) Must be displayed in a conspicuous place; and
    (2) May not be easily erased, disfigured, or obscured.
    (c) For airplanes which are to be certificated in more than one 
category--
    (1) The applicant must select one category upon which the placards 
and markings are to be based; and
    (2) The placards and marking information for all categories in which 
the airplane is to be certificated must be furnished in the Airplane 
Flight Manual.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978]



Sec. 23.1543  Instrument markings: General.

    For each instrument--
    (a) When markings are on the cover glass of the instrument, there 
must be means to maintain the correct alignment of the glass cover with 
the face of the dial; and
    (b) Each arc and line must be wide enough and located to be clearly 
visible to the pilot.
    (c) All related instruments must be calibrated in compatible units.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-50, 61 FR 5192, Feb. 9, 1996]



Sec. 23.1545  Airspeed indicator.

    (a) Each airspeed indicator must be marked as specified in paragraph 
(b) of this section, with the marks located at the corresponding 
indicated airspeeds.
    (b) The following markings must be made:
    (1) For the never-exceed speed VNE, a radial red line.
    (2) For the caution range, a yellow arc extending from the red line 
specified in paragraph (b)(1) of this section to the upper limit of the 
green arc specified in paragraph (b)(3) of this section.
    (3) For the normal operating range, a green arc with the lower limit 
at VS1 with maximum weight and with landing gear and wing flaps 
retracted, and the upper limit at the maximum structural cruising speed 
VNO established under Sec. 23.1505(b).
    (4) For the flap operating range, a white arc with the lower limit 
at VS0 at the maximum weight, and the upper limit at the flaps-extended 
speed VFE established under Sec. 23.1511.
    (5) For reciprocating multiengine-powered airplanes of 6,000 pounds 
or less maximum weight, for the speed at which compliance has been shown 
with Sec. 23.69(b) relating to rate of climb at maximum weight and at 
sea level, a blue radial line.
    (6) For reciprocating multiengine-powered airplanes of 6,000 pounds 
or less maximum weight, for the maximum value of minimum control speed, 
VMC, (one-engine-inoperative) determined under Sec. 
23.149(b), a red radial line.
    (c) If VNE or VNO vary with altitude, there must be means to 
indicate to the pilot the appropriate limitations throughout the 
operating altitude range.
    (d) Paragraphs (b)(1) through (b)(4) and paragraph (c) of this 
section do not apply to airplanes for which a maximum operating speed 
VMO/MMO is established under Sec. 23.1505(c). For 
those airplanes, there must either be a maximum allowable airspeed 
indication showing the variation of VMO/MMO with 
altitude or compressibility limitations (as appropriate), or a radial 
red line marking for VMO/MMO must be made at 
lowest value of VMO/MMO established for any 
altitude up to the maximum operating altitude for the airplane.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-3, 30 
FR 14240, Nov. 13, 1965; Amdt. 23-7, 34 FR 13097, Aug. 13, 1969; Amdt. 
23-23, 43 FR 50593, Oct. 30, 1978; Amdt. 23-50, 61 FR 5193, Feb. 9, 
1996; Amdt. 23-62, 76 FR 75762, Dec. 2, 2011]

[[Page 325]]



Sec. 23.1547  Magnetic direction indicator.

    (a) A placard meeting the requirements of this section must be 
installed on or near the magnetic direction indicator.
    (b) The placard must show the calibration of the instrument in level 
flight with the engines operating.
    (c) The placard must state whether the calibration was made with 
radio receivers on or off.
    (d) Each calibration reading must be in terms of magnetic headings 
in not more than 30 degree increments.
    (e) If a magnetic nonstabilized direction indicator can have a 
deviation of more than 10 degrees caused by the operation of electrical 
equipment, the placard must state which electrical loads, or combination 
of loads, would cause a deviation of more than 10 degrees when turned 
on.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-20, 42 FR 36969, July 18, 1977]



Sec. 23.1549  Powerplant and auxiliary power unit instruments.

    For each required powerplant and auxiliary power unit instrument, as 
appropriate to the type of instruments--
    (a) Each maximum and, if applicable, minimum safe operating limit 
must be marked with a red radial or a red line;
    (b) Each normal operating range must be marked with a green arc or 
green line, not extending beyond the maximum and minimum safe limits;
    (c) Each takeoff and precautionary range must be marked with a 
yellow arc or a yellow line; and
    (d) Each engine, auxiliary power unit, or propeller range that is 
restricted because of excessive vibration stresses must be marked with 
red arcs or red lines.

[Amdt. 23-12, 41 FR 55466, Dec. 20, 1976, as amended by Amdt. 23-28, 47 
FR 13315, Mar. 29, 1982; Amdt. 23-45, 58 FR 42166, Aug. 6, 1993]



Sec. 23.1551  Oil quantity indicator.

    Each oil quantity indicator must be marked in sufficient increments 
to indicate readily and accurately the quantity of oil.



Sec. 23.1553  Fuel quantity indicator.

    A red radial line must be marked on each indicator at the calibrated 
zero reading, as specified in Sec. 23.1337(b)(1).

[Doc. No. 27807, 61 FR 5193, Feb. 9, 1996]



Sec. 23.1555  Control markings.

    (a) Each cockpit control, other than primary flight controls and 
simple push button type starter switches, must be plainly marked as to 
its function and method of operation.
    (b) Each secondary control must be suitably marked.
    (c) For powerplant fuel controls--
    (1) Each fuel tank selector control must be marked to indicate the 
position corresponding to each tank and to each existing cross feed 
position;
    (2) If safe operation requires the use of any tanks in a specific 
sequence, that sequence must be marked on or near the selector for those 
tanks;
    (3) The conditions under which the full amount of usable fuel in any 
restricted usage fuel tank can safely be used must be stated on a 
placard adjacent to the selector valve for that tank; and
    (4) Each valve control for any engine of a multiengine airplane must 
be marked to indicate the position corresponding to each engine 
controlled.
    (d) Usable fuel capacity must be marked as follows:
    (1) For fuel systems having no selector controls, the usable fuel 
capacity of the system must be indicated at the fuel quantity indicator.
    (2) For fuel systems having selector controls, the usable fuel 
capacity available at each selector control position must be indicated 
near the selector control.
    (3) For fuel systems having a calibrated fuel quantity indication 
system complying with Sec. 23.1337(b)(1) and accurately displaying the 
actual quantity of usable fuel in each selectable tank, no fuel capacity 
placards outside of the fuel quantity indicator are required.
    (e) For accessory, auxiliary, and emergency controls--
    (1) If retractable landing gear is used, the indicator required by 
Sec. 23.729 must be marked so that the pilot can, at any time, 
ascertain that the wheels are secured in the extreme positions; and

[[Page 326]]

    (2) Each emergency control must be red and must be marked as to 
method of operation. No control other than an emergency control, or a 
control that serves an emergency function in addition to its other 
functions, shall be this color.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978; Amdt. 23-50, 61 FR 
5193, Feb. 9, 1996; Amdt. 23-62, 76 FR 75763, Dec. 2, 2011]



Sec. 23.1557  Miscellaneous markings and placards.

    (a) Baggage and cargo compartments, and ballast location. Each 
baggage and cargo compartment, and each ballast location, must have a 
placard stating any limitations on contents, including weight, that are 
necessary under the loading requirements.
    (b) Seats. If the maximum allowable weight to be carried in a seat 
is less than 170 pounds, a placard stating the lesser weight must be 
permanently attached to the seat structure.
    (c) Fuel, oil, and coolant filler openings. The following apply:
    (1) Fuel filler openings must be marked at or near the filler cover 
with--
    (i) For reciprocating engine-powered airplanes--
    (A) The word ``Avgas''; and
    (B) The minimum fuel grade.
    (ii) For turbine engine-powered airplanes--
    (A) The words ``Jet Fuel''; and
    (B) The permissible fuel designations, or references to the Airplane 
Flight Manual (AFM) for permissible fuel designations.
    (iii) For pressure fueling systems, the maximum permissible fueling 
supply pressure and the maximum permissible defueling pressure.
    (2) Oil filler openings must be marked at or near the filler cover 
with the word ``Oil'' and the permissible oil designations, or 
references to the Airplane Flight Manual (AFM) for permissible oil 
designations.
    (3) Coolant filler openings must be marked at or near the filler 
cover with the word ``Coolant''.
    (d) Emergency exit placards. Each placard and operating control for 
each emergency exit must be red. A placard must be near each emergency 
exit control and must clearly indicate the location of that exit and its 
method of operation.
    (e) The system voltage of each direct current installation must be 
clearly marked adjacent to its exernal power connection.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; as amended by Amdt. 23-21, 
42 FR 15042, Mar. 17, 1977; Amdt. 23-23, 43 FR 50594, Oct. 30, 1978; 
Amdt. 23-45, 58 FR 42166, Aug. 6, 1993; 73 FR 35063, June 20, 2008]



Sec. 23.1559  Operating limitations placard.

    (a) There must be a placard in clear view of the pilot stating--
    (1) That the airplane must be operated in accordance with the 
Airplane Flight Manual; and
    (2) The certification category of the airplane to which the placards 
apply.
    (b) For airplanes certificated in more than one category, there must 
be a placard in clear view of the pilot stating that other limitations 
are contained in the Airplane Flight Manual.
    (c) There must be a placard in clear view of the pilot that 
specifies the kind of operations to which the operation of the airplane 
is limited or from which it is prohibited under Sec. 23.1525.
    (d) The placard(s) required by this section need not be lighted.

[Doc. No. 27807, 61 FR 5193, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75763, Dec. 2, 2011]



Sec. 23.1561  Safety equipment.

    (a) Safety equipment must be plainly marked as to method of 
operation.
    (b) Stowage provisions for required safety equipment must be marked 
for the benefit of occupants.



Sec. 23.1563  Airspeed placards.

    There must be an airspeed placard in clear view of the pilot and as 
close as practicable to the airspeed indicator. This placard must list--
    (a) The operating maneuvering speed, VO; and
    (b) The maximum landing gear operating speed VLO.
    (c) For reciprocating multiengine-powered airplanes of more than 
6,000 pounds maximum weight, and turbine

[[Page 327]]

engine-powered airplanes, the maximum value of the minimum control 
speed, VMC (one-engine-inoperative) determined under Sec. 
23.149(b).
    (d) The airspeed placard(s) required by this section need not be 
lighted if the landing gear operating speed is indicated on the airspeed 
indicator or other lighted area such as the landing gear control and the 
airspeed indicator has features such as low speed awareness that provide 
ample warning prior to VMC.

[Amdt. 23-7, 34 FR 13097, Aug. 13, 1969, as amended by Amdt. 23-45, 58 
FR 42166, Aug. 6, 1993; Amdt. 23-50, 61 FR 5193, Feb. 9, 1996; Amdt. 23-
62, 76 FR 75763, Dec. 2, 2011]



Sec. 23.1567  Flight maneuver placard.

    (a) For normal category airplanes, there must be a placard in front 
of and in clear view of the pilot stating: ``No acrobatic maneuvers, 
including spins, approved.''
    (b) For utility category airplanes, there must be--
    (1) A placard in clear view of the pilot stating: ``Acrobatic 
maneuvers are limited to the following ----------------------;'' (list 
approved maneuvers and the recommended entry speed for each); and
    (2) For those airplanes that do not meet the spin requirements for 
acrobatic category airplanes, an additional placard in clear view of the 
pilot stating: ``Spins Prohibited.''
    (c) For acrobatic category airplanes, there must be a placard in 
clear view of the pilot listing the approved acrobatic maneuvers and the 
recommended entry airspeed for each. If inverted flight maneuvers are 
not approved, the placard must bear a notation to this effect.
    (d) For acrobatic category airplanes and utility category airplanes 
approved for spinning, there must be a placard in clear view of the 
pilot--
    (1) Listing the control actions for recovery from spinning 
maneuvers; and
    (2) Stating that recovery must be initiated when spiral 
characteristics appear, or after not more than six turns or not more 
than any greater number of turns for which the airplane has been 
certificated.
    (e) The placard(s) required by this section need not be lighted.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as 
amended by Amdt. 23-13, 37 FR 20023, Sept. 23, 1972; Amdt. 23-21, 43 FR 
2319, Jan. 16, 1978; Amdt. 23-50, 61 FR 5193, Feb. 9, 1996; Amdt. 23-62, 
76 FR 75763, Dec. 2, 2011]

           Airplane Flight Manual and Approved Manual Material



Sec. 23.1581  General.

    (a) Furnishing information. An Airplane Flight Manual must be 
furnished with each airplane, and it must contain the following:
    (1) Information required by Sec. Sec. 23.1583 through 23.1589.
    (2) Other information that is necessary for safe operation because 
of design, operating, or handling characteristics.
    (3) Further information necessary to comply with the relevant 
operating rules.
    (b) Approved information. (1) Except as provided in paragraph (b)(2) 
of this section, each part of the Airplane Flight Manual containing 
information prescribed in Sec. Sec. 23.1583 through 23.1589 must be 
approved, segregated, identified and clearly distinguished from each 
unapproved part of that Airplane Flight Manual.
    (2) The requirements of paragraph (b)(1) of this section do not 
apply to reciprocating engine-powered airplanes of 6,000 pounds or less 
maximum weight, if the following is met:
    (i) Each part of the Airplane Flight Manual containing information 
prescribed in Sec. 23.1583 must be limited to such information, and 
must be approved, identified, and clearly distinguished from each other 
part of the Airplane Flight Manual.
    (ii) The information prescribed in Sec. Sec. 23.1585 through 
23.1589 must be determined in accordance with the applicable 
requirements of this part and presented in its entirety in a manner 
acceptable to the Administrator.
    (3) Each page of the Airplane Flight Manual containing information 
prescribed in this section must be of a type that is not easily erased, 
disfigured, or misplaced, and is capable of being inserted in a manual 
provided by

[[Page 328]]

the applicant, or in a folder, or in any other permanent binder.
    (c) The units used in the Airplane Flight Manual must be the same as 
those marked on the appropriate instruments and placards.
    (d) All Airplane Flight Manual operational airspeeds, unless 
otherwise specified, must be presented as indicated airspeeds.
    (e) Provision must be made for stowing the Airplane Flight Manual in 
a suitable fixed container which is readily accessible to the pilot.
    (f) Revisions and amendments. Each Airplane Flight Manual (AFM) must 
contain a means for recording the incorporation of revisions and 
amendments.

[Amdt. 23-21, 43 FR 2319, Jan. 16, 1978, as amended by Amdt. 23-34, 52 
FR 1834, Jan. 15, 1987; Amdt. 23-45, 58 FR 42166, Aug. 6, 1993; Amdt. 
23-50, 61 FR 5193, Feb. 9, 1996]



Sec. 23.1583  Operating limitations.

    The Airplane Flight Manual must contain operating limitations 
determined under this part 23, including the following--
    (a) Airspeed limitations. The following information must be 
furnished:
    (1) Information necessary for the marking of the airspeed limits on 
the indicator as required in Sec. 23.1545, and the significance of each 
of those limits and of the color coding used on the indicator.
    (2) The speeds VMC, VO, VLE, and 
VLO, if established, and their significance.
    (3) In addition, for turbine powered commuter category airplanes--
    (i) The maximum operating limit speed, VMO/MMO 
and a statement that this speed must not be deliberately exceeded in any 
regime of flight (climb, cruise or descent) unless a higher speed is 
authorized for flight test or pilot training;
    (ii) If an airspeed limitation is based upon compressibility 
effects, a statement to this effect and information as to any symptoms, 
the probable behavior of the airplane, and the recommended recovery 
procedures; and
    (iii) The airspeed limits must be shown in terms of VMO/
MMO instead of VNO and VNE.
    (b) Powerplant limitations. The following information must be 
furnished:
    (1) Limitations required by Sec. 23.1521.
    (2) Explanation of the limitations, when appropriate.
    (3) Information necessary for marking the instruments required by 
Sec. 23.1549 through Sec. 23.1553.
    (c) Weight. The airplane flight manual must include--
    (1) The maximum weight; and
    (2) The maximum landing weight, if the design landing weight 
selected by the applicant is less than the maximum weight.
    (3) For reciprocating engine-powered airplanes of more than 6,000 
pounds maximum weight, single-engine turbines, and multiengine jets 
6,000 pounds or less maximum weight in the normal, utility, and 
acrobatic category, performance operating limitations as follows--
    (i) The maximum takeoff weight for each airport altitude and ambient 
temperature within the range selected by the applicant at which the 
airplane complies with the climb requirements of Sec. 23.63(c)(1).
    (ii) The maximum landing weight for each airport altitude and 
ambient temperature within the range selected by the applicant at which 
the airplane complies with the climb requirements of Sec. 23.63(c)(2).
    (4) For normal, utility, and acrobatic category multiengine jets 
over 6,000 pounds and commuter category airplanes, the maximum takeoff 
weight for each airport altitude and ambient temperature within the 
range selected by the applicant at which--
    (i) The airplane complies with the climb requirements of Sec. 
23.63(d)(1); and
    (ii) The accelerate-stop distance determined under Sec. 23.55 is 
equal to the available runway length plus the length of any stopway, if 
utilized; and either:
    (A) The takeoff distance determined under Sec. 23.59(a) is equal to 
the available runway length; or
    (B) At the option of the applicant, the takeoff distance determined 
under Sec. 23.59(a) is equal to the available runway length plus the 
length of any clearway and the takeoff run determined under Sec. 
23.59(b) is equal to the available runway length.
    (5) For normal, utility, and acrobatic category multiengine jets 
over 6,000

[[Page 329]]

pounds and commuter category airplanes, the maximum landing weight for 
each airport altitude within the range selected by the applicant at 
which--
    (i) The airplane complies with the climb requirements of Sec. 
23.63(d)(2) for ambient temperatures within the range selected by the 
applicant; and
    (ii) The landing distance determined under Sec. 23.75 for standard 
temperatures is equal to the available runway length.
    (6) The maximum zero wing fuel weight, where relevant, as 
established in accordance with Sec. 23.343.
    (d) Center of gravity. The established center of gravity limits.
    (e) Maneuvers. The following authorized maneuvers, appropriate 
airspeed limitations, and unauthorized maneuvers, as prescribed in this 
section.
    (1) Normal category airplanes. No acrobatic maneuvers, including 
spins, are authorized.
    (2) Utility category airplanes. A list of authorized maneuvers 
demonstrated in the type flight tests, together with recommended entry 
speeds and any other associated limitations. No other maneuver is 
authorized.
    (3) Acrobatic category airplanes. A list of approved flight 
maneuvers demonstrated in the type flight tests, together with 
recommended entry speeds and any other associated limitations.
    (4) Acrobatic category airplanes and utility category airplanes 
approved for spinning. Spin recovery procedure established to show 
compliance with Sec. 23.221(c).
    (5) Commuter category airplanes. Maneuvers are limited to any 
maneuver incident to normal flying, stalls, (except whip stalls) and 
steep turns in which the angle of bank is not more than 60 degrees.
    (f) Maneuver load factor. The positive limit load factors in g's, 
and, in addition, the negative limit load factor for acrobatic category 
airplanes.
    (g) Minimum flight crew. The number and functions of the minimum 
flight crew determined under Sec. 23.1523.
    (h) Kinds of operation. A list of the kinds of operation to which 
the airplane is limited or from which it is prohibited under Sec. 
23.1525, and also a list of installed equipment that affects any 
operating limitation and identification as to the equipment's required 
operational status for the kinds of operation for which approval has 
been given.
    (i) Maximum operating altitude. The maximum altitude established 
under Sec. 23.1527.
    (j) Maximum passenger seating configuration. The maximum passenger 
seating configuration.
    (k) Allowable lateral fuel loading. The maximum allowable lateral 
fuel loading differential, if less than the maximum possible.
    (l) Baggage and cargo loading. The following information for each 
baggage and cargo compartment or zone--
    (1) The maximum allowable load; and
    (2) The maximum intensity of loading.
    (m) Systems. Any limitations on the use of airplane systems and 
equipment.
    (n) Ambient temperatures. Where appropriate, maximum and minimum 
ambient air temperatures for operation.
    (o) Smoking. Any restrictions on smoking in the airplane.
    (p) Types of surface. A statement of the types of surface on which 
operations may be conducted. (See Sec. 23.45(g) and Sec. 23.1587 
(a)(4), (c)(2), and (d)(4)).

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13097, Aug. 13, 1969; Amdt. 23-10, 36 FR 2864, Feb. 11, 1971; Amdt. 
23-21, 43 FR 2320, Jan. 16, 1978; Amdt. 23-23, 43 FR 50594, Oct. 30, 
1978; Amdt. 23-34, 52 FR 1834, Jan. 15, 1987; Amdt. 23-45, 58 FR 42166, 
Aug. 6, 1993; Amdt. 23-50, 61 FR 5193, Feb. 9, 1996; Amdt. 23-62, 76 FR 
75763, Dec. 2, 2011]



Sec. 23.1585  Operating procedures.

    (a) For all airplanes, information concerning normal, abnormal (if 
applicable), and emergency procedures and other pertinent information 
necessary for safe operation and the achievement of the scheduled 
performance must be furnished, including--
    (1) An explanation of significant or unusual flight or ground 
handling characteristics;
    (2) The maximum demonstrated values of crosswind for takeoff and 
landing, and procedures and information pertinent to operations in 
crosswinds;
    (3) A recommended speed for flight in rough air. This speed must be 
chosen to

[[Page 330]]

protect against the occurrence, as a result of gusts, of structural 
damage to the airplane and loss of control (for example, stalling);
    (4) Procedures for restarting any turbine engine in flight, 
including the effects of altitude; and
    (5) Procedures, speeds, and configuration(s) for making a normal 
approach and landing, in accordance with Sec. Sec. 23.73 and 23.75, and 
a transition to the balked landing condition.
    (6) For seaplanes and amphibians, water handling procedures and the 
demonstrated wave height.
    (b) In addition to paragraph (a) of this section, for all single-
engine airplanes, the procedures, speeds, and configuration(s) for a 
glide following engine failure, in accordance with Sec. 23.71 and the 
subsequent forced landing, must be furnished.
    (c) In addition to paragraph (a) of this section, for all 
multiengine airplanes, the following information must be furnished:
    (1) Procedures, speeds, and configuration(s) for making an approach 
and landing with one engine inoperative;
    (2) Procedures, speeds, and configuration(s) for making a balked 
landing with one engine inoperative and the conditions under which a 
balked landing can be performed safely, or a warning against attempting 
a balked landing;
    (3) The VSSE determined in Sec. 23.149; and
    (4) Procedures for restarting any engine in flight including the 
effects of altitude.
    (d) In addition to paragraphs (a) and either (b) or (c) of this 
section, as appropriate, for all normal, utility, and acrobatic category 
airplanes, the following information must be furnished:
    (1) Procedures, speeds, and configuration(s) for making a normal 
takeoff, in accordance with Sec. 23.51 (a) and (b), and Sec. 23.53 (a) 
and (b), and the subsequent climb, in accordance with Sec. 23.65 and 
Sec. 23.69(a).
    (2) Procedures for abandoning a takeoff due to engine failure or 
other cause.
    (e) In addition to paragraphs (a), (c), and (d) of this section, for 
all normal, utility, and acrobatic category multiengine airplanes, the 
information must include the following:
    (1) Procedures and speeds for continuing a takeoff following engine 
failure and the conditions under which takeoff can safely be continued, 
or a warning against attempting to continue the takeoff.
    (2) Procedures, speeds, and configurations for continuing a climb 
following engine failure, after takeoff, in accordance with Sec. 23.67, 
or enroute, in accordance with Sec. 23.69(b).
    (f) In addition to paragraphs (a) and (c) of this section, for 
normal, utility, and acrobatic category multiengine jets weighing over 
6,000 pounds, and commuter category airplanes, the information must 
include the following:
    (1) Procedures, speeds, and configuration(s) for making a normal 
takeoff.
    (2) Procedures and speeds for carrying out an accelerate-stop in 
accordance with Sec. 23.55.
    (3) Procedures and speeds for continuing a takeoff following engine 
failure in accordance with Sec. 23.59(a)(1) and for following the 
flight path determined under Sec. 23.57 and Sec. 23.61(a).
    (g) For multiengine airplanes, information identifying each 
operating condition in which the fuel system independence prescribed in 
Sec. 23.953 is necessary for safety must be furnished, together with 
instructions for placing the fuel system in a configuration used to show 
compliance with that section.
    (h) For each airplane showing compliance with Sec. 23.1353 (g)(2) 
or (g)(3), the operating procedures for disconnecting the battery from 
its charging source must be furnished.
    (i) Information on the total quantity of usable fuel for each fuel 
tank, and the effect on the usable fuel quantity, as a result of a 
failure of any pump, must be furnished.
    (j) Procedures for the safe operation of the airplane's systems and 
equipment, both in normal use and in the event of malfunction, must be 
furnished.

[Doc. No. 27807, 61 FR 5194, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75763, Dec. 2, 2011]



Sec. 23.1587  Performance information.

    Unless otherwise prescribed, performance information must be 
provided over the altitude and temperature ranges required by Sec. 
23.45(b).

[[Page 331]]

    (a) For all airplanes, the following information must be furnished--
    (1) The stalling speeds VSO and VS1 with the 
landing gear and wing flaps retracted, determined at maximum weight 
under Sec. 23.49, and the effect on these stalling speeds of angles of 
bank up to 60 degrees;
    (2) The steady rate and gradient of climb with all engines 
operating, determined under Sec. 23.69(a);
    (3) The landing distance, determined under Sec. 23.75 for each 
airport altitude and standard temperature, and the type of surface for 
which it is valid;
    (4) The effect on landing distances of operation on other than 
smooth hard surfaces, when dry, determined under Sec. 23.45(g); and
    (5) The effect on landing distances of runway slope and 50 percent 
of the headwind component and 150 percent of the tailwind component.
    (b) In addition to paragraph (a) of this section, for all normal, 
utility, and acrobatic category reciprocating engine-powered airplanes 
of 6,000 pounds or less maximum weight, the steady angle of climb/
descent, determined under Sec. 23.77(a), must be furnished.
    (c) In addition to paragraphs (a) and (b) of this section, if 
appropriate, for normal, utility, and acrobatic category airplanes, the 
following information must be furnished--
    (1) The takeoff distance, determined under Sec. 23.53 and the type 
of surface for which it is valid.
    (2) The effect on takeoff distance of operation on other than smooth 
hard surfaces, when dry, determined under Sec. 23.45(g);
    (3) The effect on takeoff distance of runway slope and 50 percent of 
the headwind component and 150 percent of the tailwind component;
    (4) For multiengine reciprocating engine-powered airplanes of more 
than 6,000 pounds maximum weight and multiengine turbine powered 
airplanes, the one-engine-inoperative takeoff climb/descent gradient, 
determined under Sec. 23.66;
    (5) For multiengine airplanes, the enroute rate and gradient of 
climb/descent with one engine inoperative, determined under Sec. 
23.69(b); and
    (6) For single-engine airplanes, the glide performance determined 
under Sec. 23.71.
    (d) In addition to paragraph (a) of this section, for normal, 
utility, and acrobatic category multiengine jets weighing over 6,000 
pounds, and commuter category airplanes, the following information must 
be furnished--
    (1) The accelerate-stop distance determined under Sec. 23.55;
    (2) The takeoff distance determined under Sec. 23.59(a);
    (3) At the option of the applicant, the takeoff run determined under 
Sec. 23.59(b);
    (4) The effect on accelerate-stop distance, takeoff distance and, if 
determined, takeoff run, of operation on other than smooth hard 
surfaces, when dry, determined under Sec. 23.45(g);
    (5) The effect on accelerate-stop distance, takeoff distance, and if 
determined, takeoff run, of runway slope and 50 percent of the headwind 
component and 150 percent of the tailwind component;
    (6) The net takeoff flight path determined under Sec. 23.61(b);
    (7) The enroute gradient of climb/descent with one engine 
inoperative, determined under Sec. 23.69(b);
    (8) The effect, on the net takeoff flight path and on the enroute 
gradient of climb/descent with one engine inoperative, of 50 percent of 
the headwind component and 150 percent of the tailwind component;
    (9) Overweight landing performance information (determined by 
extrapolation and computed for the range of weights between the maximum 
landing and maximum takeoff weights) as follows--
    (i) The maximum weight for each airport altitude and ambient 
temperature at which the airplane complies with the climb requirements 
of Sec. 23.63(d)(2); and
    (ii) The landing distance determined under Sec. 23.75 for each 
airport altitude and standard temperature.
    (10) The relationship between IAS and CAS determined in accordance 
with Sec. 23.1323 (b) and (c).

[[Page 332]]

    (11) The altimeter system calibration required by Sec. 23.1325(e).

[Doc. No. 27807, 61 FR 5194, Feb. 9, 1996, as amended by Amdt. 23-62, 76 
FR 75763, Dec. 2, 2011]



Sec. 23.1589  Loading information.

    The following loading information must be furnished:
    (a) The weight and location of each item of equipment that can be 
easily removed, relocated, or replaced and that is installed when the 
airplane was weighed under the requirement of Sec. 23.25.
    (b) Appropriate loading instructions for each possible loading 
condition between the maximum and minimum weights established under 
Sec. 23.25, to facilitate the center of gravity remaining within the 
limits established under Sec. 23.23.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42167, Aug. 6, 1993; Amdt. 23-50, 61 FR 5195, Feb. 9, 1996]



       Sec. Appendix A to Part 23--Simplified Design Load Criteria

                             A23.1 General.

    (a) The design load criteria in this appendix are an approved 
equivalent of those in Sec. Sec. 23.321 through 23.459 of this 
subchapter for an airplane having a maximum weight of 6,000 pounds or 
less and the following configuration:
    (1) A single engine excluding turbine powerplants;
    (2) A main wing located closer to the airplane's center of gravity 
than to the aft, fuselage-mounted, empennage;
    (3) A main wing that contains a quarter-chord sweep angle of not 
more than 15 degrees fore or aft;
    (4) A main wing that is equipped with trailing-edge controls 
(ailerons or flaps, or both);
    (5) A main wing aspect ratio not greater than 7;
    (6) A horizontal tail aspect ratio not greater than 4;
    (7) A horizontal tail volume coefficient not less than 0.34;
    (8) A vertical tail aspect ratio not greater than 2;
    (9) A vertical tail platform area not greater than 10 percent of the 
wing platform area; and
    (10) Symmetrical airfoils must be used in both the horizontal and 
vertical tail designs.
    (b) Appendix A criteria may not be used on any airplane 
configuration that contains any of the following design features:
    (1) Canard, tandem-wing, close-coupled, or tailless arrangements of 
the lifting surfaces;
    (2) Biplane or multiplane wing arrangements;
    (3) T-tail, V-tail, or cruciform-tail (+) arrangements;
    (4) Highly-swept wing platform (more than 15-degrees of sweep at the 
quarter-chord), delta planforms, or slatted lifting surfaces; or
    (5) Winglets or other wing tip devices, or outboard fins.

                         A23.3 Special symbols.

n1=Airplane Positive Maneuvering Limit Load Factor.
n2=Airplane Negative Maneuvering Limit Load Factor.
n3=Airplane Positive Gust Limit Load Factor at VC.
n4=Airplane Negative Gust Limit Load Factor at VC.
nflap=Airplane Positive Limit Load Factor With Flaps Fully Extended at 
          VF.
          [GRAPHIC] [TIFF OMITTED] TC28SE91.020
          
             A23.5 Certification in more than one category.

    The criteria in this appendix may be used for certification in the 
normal, utility, and acrobatic categories, or in any combination of 
these categories. If certification in more than one category is desired, 
the design category weights must be selected to make the term n1W 
constant for all categories or greater for one desired category than for 
others. The wings and control surfaces (including wing flaps and tabs) 
need only be investigated for the maximum value of n1W, or for the 
category corresponding to the maximum design weight, where n1W is 
constant. If the acrobatic category is selected, a special unsymmetrical 
flight load investigation in accordance with paragraphs A23.9(c)(2) and 
A23.11(c)(2) of this appendix must be completed. The wing, wing 
carrythrough, and the horizontal tail structures must be checked for 
this condition. The basic fuselage structure need only be investigated 
for the highest load factor design category selected. The local 
supporting structure for dead weight items need only be designed for the 
highest load factor imposed when the

[[Page 333]]

particular items are installed in the airplane. The engine mount, 
however, must be designed for a higher side load factor, if 
certification in the acrobatic category is desired, than that required 
for certification in the normal and utility categories. When designing 
for landing loads, the landing gear and the airplane as a whole need 
only be investigated for the category corresponding to the maximum 
design weight. These simplifications apply to single-engine aircraft of 
conventional types for which experience is available, and the 
Administrator may require additional investigations for aircraft with 
unusual design features.

                           A23.7 Flight loads.

    (a) Each flight load may be considered independent of altitude and, 
except for the local supporting structure for dead weight items, only 
the maximum design weight conditions must be investigated.
    (b) Table 1 and figures 3 and 4 of this appendix must be used to 
determine values of n1, n2, n3, and n4, corresponding to the maximum 
design weights in the desired categories.
    (c) Figures 1 and 2 of this appendix must be used to determine 
values of n3 and n4 corresponding to the minimum flying weights in the 
desired categories, and, if these load factors are greater than the load 
factors at the design weight, the supporting structure for dead weight 
items must be substantiated for the resulting higher load factors.
    (d) Each specified wing and tail loading is independent of the 
center of gravity range. The applicant, however, must select a c.g. 
range, and the basic fuselage structure must be investigated for the 
most adverse dead weight loading conditions for the c.g. range selected.
    (e) The following loads and loading conditions are the minimums for 
which strength must be provided in the structure:
    (1) Airplane equilibrium. The aerodynamic wing loads may be 
considered to act normal to the relative wind, and to have a magnitude 
of 1.05 times the airplane normal loads (as determined from paragraphs 
A23.9 (b) and (c) of this appendix) for the positive flight conditions 
and a magnitude equal to the airplane normal loads for the negative 
conditions. Each chordwise and normal component of this wing load must 
be considered.
    (2) Minimum design airspeeds. The minimum design airspeeds may be 
chosen by the applicant except that they may not be less than the 
minimum speeds found by using figure 3 of this appendix. In addition, 
VCmin need not exceed values of 0.9 VH actually obtained at sea level 
for the lowest design weight category for which certification is 
desired. In computing these minimum design airspeeds, n1 may not be less 
than 3.8.
    (3) Flight load factor. The limit flight load factors specified in 
Table 1 of this appendix represent the ratio of the aerodynamic force 
component (acting normal to the assumed longitudinal axis of the 
airplane) to the weight of the airplane. A positive flight load factor 
is an aerodynamic force acting upward, with respect to the airplane.

                        A23.9 Flight conditions.

    (a) General. Each design condition in paragraphs (b) and (c) of this 
section must be used to assure sufficient strength for each condition of 
speed and load factor on or within the boundary of a V-n diagram for the 
airplane similar to the diagram in figure 4 of this appendix. This 
diagram must also be used to determine the airplane structural operating 
limitations as specified in Sec. Sec. 23.1501(c) through 23.1513 and 
Sec. 23.1519.
    (b) Symmetrical flight conditions. The airplane must be designed for 
symmetrical flight conditions as follows:
    (1) The airplane must be designed for at least the four basic flight 
conditions, ``A'', ``D'', ``E'', and ``G'' as noted on the flight 
envelope of figure 4 of this appendix. In addition, the following 
requirements apply:
    (i) The design limit flight load factors corresponding to conditions 
``D'' and ``E'' of figure 4 must be at least as great as those specified 
in Table 1 and figure 4 of this appendix, and the design speed for these 
conditions must be at least equal to the value of VD found from figure 3 
of this appendix.
    (ii) For conditions ``A'' and ``G'' of figure 4, the load factors 
must correspond to those specified in Table 1 of this appendix, and the 
design speeds must be computed using these load factors with the maximum 
static lift coefficient CNA determined by the applicant. However, in the 
absence of more precise computations, these latter conditions may be 
based on a value of CNA=1.35 and the design speed 
for condition ``A'' may be less than VAmin.
    (iii) Conditions ``C'' and ``F'' of figure 4 need only be 
investigated when n3 W/S or n4 W/S are greater than n1 W/S or n2 W/S of 
this appendix, respectively.
    (2) If flaps or other high lift devices intended for use at the 
relatively low airspeed of approach, landing, and takeoff, are 
installed, the airplane must be designed for the two flight conditions 
corresponding to the values of limit flap-down factors specified in 
Table 1 of this appendix with the flaps fully extended at not less than 
the design flap speed VFmin from figure 3 of this appendix.
    (c) Unsymmetrical flight conditions. Each affected structure must be 
designed for unsymmetrical loadings as follows:
    (1) The aft fuselage-to-wing attachment must be designed for the 
critical vertical surface load determined in accordance with paragraph 
SA23.11(c)(1) and (2) of this appendix.

[[Page 334]]

    (2) The wing and wing carry-through structures must be designed for 
100 percent of condition ``A'' loading on one side of the plane of 
symmetry and 70 percent on the opposite side for certification in the 
normal and utility categories, or 60 percent on the opposite side for 
certification in the acrobatic category.
    (3) The wing and wing carry-through structures must be designed for 
the loads resulting from a combination of 75 percent of the positive 
maneuvering wing loading on both sides of the plane of symmetry and the 
maximum wing torsion resulting from aileron displacement. The effect of 
aileron displacement on wing torsion at VC or VA using the basic airfoil 
moment coefficient modified over the aileron portion of the span, must 
be computed as follows:
    (i) Cm=Cm +0.01[delta][mu] (up aileron side) wing basic 
airfoil.
    (ii) Cm=Cm -0.01[delta][mu](down aileron side) wing basic 
airfoil, where [delta][mu] is the up aileron deflection and 
[delta] d is the down aileron deflection.
    (4) [Delta] critical, which is the sum of 
[delta][mu]+[delta] d must be computed as follows:
    (i) Compute [Delta][alpha] and [Delta]b from the formulas:
    [GRAPHIC] [TIFF OMITTED] TC28SE91.021
    
Where [Delta]p=the maximum total deflection (sum of both aileron 
          deflections) at VA with VA, VC, and VD described in 
          subparagraph (2) of Sec. 23.7(e) of this appendix.

    (ii) Compute K from the formula:
    [GRAPHIC] [TIFF OMITTED] TC28SE91.022
    
where [delta][alpha] is the down aileron deflection 
          corresponding to [Delta][alpha], and 
          [delta]b is the down aileron deflection 
          corresponding to [Delta] b as computed in step (i).
    (iii) If K is less than 1.0, [Delta][alpha] is [Delta] 
critical and must be used to determine [delta]u and 
[delta]d. In this case, VC is the critical speed which must 
be used in computing the wing torsion loads over the aileron span.
    (iv) If K is equal to or greater than 1.0, [Delta]b is 
[Delta] critical and must be used to determine [delta]u and 
[delta]d. In this case, Vd is the critical speed which must 
be used in computing the wing torsion loads over the aileron span.
    (d) Supplementary conditions; rear lift truss; engine torque; side 
load on engine mount. Each of the following supplementary conditions 
must be investigated:
    (1) In designing the rear lift truss, the special condition 
specified in Sec. 23.369 may be investigated instead of condition ``G'' 
of figure 4 of this appendix. If this is done, and if certification in 
more than one category is desired, the value of W/S used in the formula 
appearing in Sec. 23.369 must be that for the category corresponding to 
the maximum gross weight.
    (2) Each engine mount and its supporting structures must be designed 
for the maximum limit torque corresponding to METO power and propeller 
speed acting simultaneously with the limit loads resulting from the 
maximum positive maneuvering flight load factor n1. The limit torque 
must be obtained by multiplying the mean torque by a factor of 1.33 for 
engines with five or more cylinders. For 4, 3, and 2 cylinder engines, 
the factor must be 2, 3, and 4, respectively.
    (3) Each engine mount and its supporting structure must be designed 
for the loads resulting from a lateral limit load factor of not less 
than 1.47 for the normal and utility categories, or 2.0 for the 
acrobatic category.

                      A23.11 Control surface loads.

    (a) General. Each control surface load must be determined using the 
criteria of paragraph (b) of this section and must lie within the 
simplified loadings of paragraph (c) of this section.
    (b) Limit pilot forces. In each control surface loading condition 
described in paragraphs (c) through (e) of this section, the airloads on 
the movable surfaces and the corresponding deflections need not exceed 
those which could be obtained in flight by employing the maximum limit 
pilot forces specified in the table in Sec. 23.397(b). If the surface 
loads are limited by these maximum limit pilot forces, the tabs must 
either be considered to be deflected to their maximum travel in the 
direction which would assist the pilot or the deflection must correspond 
to the maximum degree of ``out of trim'' expected at the speed for the 
condition under consideration. The tab load, however, need not exceed 
the value specified in Table 2 of this appendix.
    (c) Surface loading conditions. Each surface loading condition must 
be investigated as follows:
    (1) Simplified limit surface loadings for the horizontal tail, 
vertical tail, aileron, wing flaps, and trim tabs are specified in 
figures 5 and 6 of this appendix.
    (i) The distribution of load along the span of the surface, 
irrespective of the chordwise load distribution, must be assumed 
proportional to the total chord, except on horn balance surfaces.
    (ii) The load on the stabilizer and elevator, and the load on fin 
and rudder, must be distributed chordwise as shown in figure 7 of this 
appendix.
    (iii) In order to ensure adequate torsional strength and to account 
for maneuvers and

[[Page 335]]

gusts, the most severe loads must be considered in association with 
every center of pressure position between the leading edge and the half 
chord of the mean chord of the surface (stabilizer and elevator, or fin 
and rudder).
    (iv) To ensure adequate strength under high leading edge loads, the 
most severe stabilizer and fin loads must be further considered as being 
increased by 50 percent over the leading 10 percent of the chord with 
the loads aft of this appropriately decreased to retain the same total 
load.
    (v) The most severe elevator and rudder loads should be further 
considered as being distributed parabolically from three times the mean 
loading of the surface (stabilizer and elevator, or fin and rudder) at 
the leading edge of the elevator and rudder, respectively, to zero at 
the trailing edge according to the equation:
[GRAPHIC] [TIFF OMITTED] TR09FE96.004

[GRAPHIC] [TIFF OMITTED] TR09FE96.007

Where--

P(x)=local pressure at the chordwise stations x,
c=chord length of the tail surface,
cf=chord length of the elevator and rudder respectively, and
w=average surface loading as specified in Figure A5.

    (vi) The chordwise loading distribution for ailerons, wing flaps, 
and trim tabs are specified in Table 2 of this appendix.
    (2) If certification in the acrobatic category is desired, the 
horizontal tail must be investigated for an unsymmetrical load of 100 
percent w on one side of the airplane centerline and 50 percent on the 
other side of the airplane centerline.
    (d) Outboard fins. Outboard fins must meet the requirements of Sec. 
23.445.
    (e) Special devices. Special devices must meet the requirements of 
Sec. 23.459.

                      A23.13 Control system loads.

    (a) Primary flight controls and systems. Each primary flight control 
and system must be designed as follows:
    (1) The flight control system and its supporting structure must be 
designed for loads corresponding to 125 percent of the computed hinge 
moments of the movable control surface in the conditions prescribed in 
A23.11 of this appendix. In addition--
    (i) The system limit loads need not exceed those that could be 
produced by the pilot and automatic devices operating the controls; and
    (ii) The design must provide a rugged system for service use, 
including jamming, ground gusts, taxiing downwind, control inertia, and 
friction.
    (2) Acceptable maximum and minimum limit pilot forces for elevator, 
aileron, and rudder controls are shown in the table in Sec. 23.397(b). 
These pilots loads must be assumed to act at the appropriate control 
grips

[[Page 336]]

or pads as they would under flight conditions, and to be reacted at the 
attachments of the control system to the control surface horn.
    (b) Dual controls. If there are dual controls, the systems must be 
designed for pilots operating in opposition, using individual pilot 
loads equal to 75 percent of those obtained in accordance with paragraph 
(a) of this section, except that individual pilot loads may not be less 
than the minimum limit pilot forces shown in the table in Sec. 
23.397(b).
    (c) Ground gust conditions. Ground gust conditions must meet the 
requirements of Sec. 23.415.
    (d) Secondary controls and systems. Secondary controls and systems 
must meet the requirements of Sec. 23.405.

                   Table 1--Limit Flight Load Factors
                       [Limit flight load factors]
------------------------------------------------------------------------
                                        Normal      Utility    Acrobatic
         Flight load factors           category    category    category
------------------------------------------------------------------------
Flaps up:
  n1................................         3.8         4.4         6.0
  n2................................     -0.5 n1  ..........  ..........
  n3................................       (\1\)  ..........  ..........
  n4................................       (\2\)  ..........  ..........
Flaps down:
  n flap............................      0.5 n1  ..........  ..........
  n flap............................    \3\ Zero  ..........  ..........
------------------------------------------------------------------------
\1\ Find n3 from Fig. 1
\2\ Find n4 from Fig. 2
\3\ Vertical wing load may be assumed equal to zero and only the flap
  part of the wing need be checked for this condition.

  [GRAPHIC] [TIFF OMITTED] TR09FE96.008
  

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[GRAPHIC] [TIFF OMITTED] TC28SE91.023


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[GRAPHIC] [TIFF OMITTED] TC28SE91.024


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[GRAPHIC] [TIFF OMITTED] TC28SE91.025

[GRAPHIC] [TIFF OMITTED] TC28SE91.026


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[GRAPHIC] [TIFF OMITTED] TC28SE91.027


[[Page 341]]

[GRAPHIC] [TIFF OMITTED] TC28SE91.028


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 Figure A7--Chordwise Load Distribution for Stabilizer and Elevator or 
                             Fin and Rudder
[GRAPHIC] [TIFF OMITTED] TR09FE96.009

[GRAPHIC] [TIFF OMITTED] TR09FE96.005

where:

w=average surface loading (as specified in figure A.5)
E=ratio of elevator (or rudder) chord to total stabilizer and elevator 
          (or fin and rudder) chord.
d'=ratio of distance of center of pressure of a unit spanwise length of 
          combined stabilizer and elevator (or fin and rudder) measured 
          from stabilizer (or fin) leading edge to the local chord. Sign 
          convention is positive when center of pressure is behind 
          leading edge.
c=local chord.

    Note: Positive values of w, P1 and P2 are all 
measured in the same direction.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13097, Aug. 13, 1969; 34 FR 14727, Sept. 24, 1969; Amdt. 23-16, 40 FR 
2577, Jan. 14, 1975; Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. 23-
48, 61 FR 5149, Feb. 9, 1996]



                  Sec. Appendix B to Part 23 [Reserved]

             Appendix C to Part 23--Basic Landing Conditions

                                                            [C23.1 Basic landing conditions]
--------------------------------------------------------------------------------------------------------------------------------------------------------
                                                     Tail wheel type                                           Nose wheel type
                                    --------------------------------------------------------------------------------------------------------------------
             Condition                                                                                        Level landing with
                                          Level landing         Tail-down landing      Level landing with   nose wheel just clear    Tail-down landing
                                                                                       inclined reactions         of ground
--------------------------------------------------------------------------------------------------------------------------------------------------------
Reference section..................  23.479(a)(1)..........  23.481(a)(1)..........  23.479(a)(2)(i)......  23.479(a)(2)(ii).....  23.481(a)(2) and (b).
--------------------------------------------------------------------------------------------------------------------------------------------------------
Vertical component at c. g.........  nW....................  nW....................  nW...................  nW...................  nW.
Fore and aft component at c. g.....  KnW...................  0.....................  KnW..................  KnW..................  0.
Lateral component in either          0.....................  0.....................  0....................  0....................  0.
 direction at c. g.
Shock absorber extension (hydraulic  Note (2)..............  Note (2)..............  Note (2).............  Note (2).............  Note (2).
 shock absorber).
Shock absorber deflection (rubber    100...................  100...................  100..................  100..................  100.
 or spring shock absorber), percent.
Tire deflection....................  Static................  Static................  Static...............  Static...............  Static.
Main wheel loads (both wheels) (Vr)  (n-L)W................  (n-L)W b/d............  (n-L)W a'/d'.........  (n-L)W...............  (n-L)W.
Main wheel loads (both wheels) (Dr)  KnW...................  0.....................  KnW a'/d'............  KnW..................  0.

[[Page 343]]

 
Tail (nose) wheel loads (Vf).......  0.....................  (n-L)W a/d............  (n-L)W b'/d'.........  0....................  0.
Tail (nose) wheel loads (Df).......  0.....................  0.....................  KnW b'/d'............  0....................  0.
Notes..............................  (1), (3), and (4).....  (4)...................  (1)..................  (1), (3), and (4)....  (3) and (4).
--------------------------------------------------------------------------------------------------------------------------------------------------------
Note (1). K may be determined as follows: K=0.25 for W=3,000 pounds or less; K=0.33 for W=6,000 pounds or greater, with linear variation of K between
  these weights.
Note (2). For the purpose of design, the maximum load factor is assumed to occur throughout the shock absorber stroke from 25 percent deflection to 100
  percent deflection unless otherwise shown and the load factor must be used with whatever shock absorber extension is most critical for each element of
  the landing gear.
Note (3). Unbalanced moments must be balanced by a rational or conservative method.
Note (4). L is defined in Sec. 23.725(b).
Note (5). n is the limit inertia load factor, at the c.g. of the airplane, selected under Sec. 23.473 (d), (f), and (g).

[GRAPHIC] [TIFF OMITTED] TC28SE91.029


[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 
FR 13099, Aug. 13, 1969]



     Sec. Appendix D to Part 23--Wheel Spin-Up and Spring-Back Loads

D23.1 Wheel spin-up loads.
    (a) The following method for determining wheel spin-up loads for 
landing conditions is based on NACA T.N. 863. However, the drag 
component used for design may not be less than the drag load prescribed 
in Sec. 23.479(b).

FHmax=1/re [radic] 2Iw(VH--Vc)nFVmax/tS
where--

FHmax=maximum rearward horizontal force acting on the wheel (in pounds);
re=effective rolling radius of wheel under impact based on recommended 
          operating tire pressure (which may be assumed to be equal to 
          the rolling radius under a static load of njWe) in feet;

[[Page 344]]

Iw=rotational mass moment of inertia of rolling assembly (in slug feet);
VH=linear velocity of airplane parallel to ground at instant of contact 
          (assumed to be 1.2 VS0, in feet per second);
Vc=peripheral speed of tire, if prerotation is used (in feet per second) 
          (there must be a positive means of pre-rotation before pre-
          rotation may be considered);
n=equals effective coefficient of friction (0.80 may be used);

FVmax=maximum vertical force on wheel (pounds)=njWe, where We and nj are 
          defined in Sec. 23.725;

ts=time interval between ground contact and attainment of maximum 
          vertical force on wheel (seconds). (However, if the value of 
          FVmax, from the above equation exceeds 0.8 FVmax, the latter 
          value must be used for FHmax.)

    (b) The equation assumes a linear variation of load factor with time 
until the peak load is reached and under this assumption, the equation 
determines the drag force at the time that the wheel peripheral velocity 
at radius re equals the airplane velocity. Most shock absorbers do not 
exactly follow a linear variation of load factor with time. Therefore, 
rational or conservative allowances must be made to compensate for these 
variations. On most landing gears, the time for wheel spin-up will be 
less than the time required to develop maximum vertical load factor for 
the specified rate of descent and forward velocity. For exceptionally 
large wheels, a wheel peripheral velocity equal to the ground speed may 
not have been attained at the time of maximum vertical gear load. 
However, as stated above, the drag spin-up load need not exceed 0.8 of 
the maximum vertical loads.
    (c) Dynamic spring-back of the landing gear and adjacent structure 
at the instant just after the wheels come up to speed may result in 
dynamic forward acting loads of considerable magnitude. This effect must 
be determined, in the level landing condition, by assuming that the 
wheel spin-up loads calculated by the methods of this appendix are 
reversed. Dynamic spring-back is likely to become critical for landing 
gear units having wheels of large mass or high landing speeds.

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-45, 
58 FR 42167, Aug. 6, 1993]



                  Sec. Appendix E to Part 23 [Reserved]



               Sec. Appendix F to Part 23--Test Procedure

 Part I--Acceptable Test Procedure for Self-Extinguishing Materials for 
     Showing Compliance With Sec. Sec. 23.853, 23.855, and 23.1359

    Acceptable test procedure for self-extinguishing materials for 
showing compliance with Sec. Sec. 23.853, 23.855 and 23.1359.
    (a) Conditioning. Specimens must be conditioned to 70 degrees F, 
plus or minus 5 degrees, and at 50 percent plus or minus 5 percent 
relative humidity until moisture equilibrium is reached or for 24 hours. 
Only one specimen at a time may be removed from the conditioning 
environment immediately before subjecting it to the flame.
    (b) Specimen configuration. Except as provided for materials used in 
electrical wire and cable insulation and in small parts, materials must 
be tested either as a section cut from a fabricated part as installed in 
the airplane or as a specimen simulating a cut section, such as: a 
specimen cut from a flat sheet of the material or a model of the 
fabricated part. The specimen may be cut from any location in a 
fabricated part; however, fabricated units, such as sandwich panels, may 
not be separated for a test. The specimen thickness must be no thicker 
than the minimum thickness to be qualified for use in the airplane, 
except that: (1) Thick foam parts, such as seat cushions, must be tested 
in \1/2\ inch thickness; (2) when showing compliance with Sec. 
23.853(d)(3)(v) for materials used in small parts that must be tested, 
the materials must be tested in no more than \1/8\ inch thickness; (3) 
when showing compliance with Sec. 23.1359(c) for materials used in 
electrical wire and cable insulation, the wire and cable specimens must 
be the same size as used in the airplane. In the case of fabrics, both 
the warp and fill direction of the weave must be tested to determine the 
most critical flammability conditions. When performing the tests 
prescribed in paragraphs (d) and (e) of this appendix, the specimen must 
be mounted in a metal frame so that (1) in the vertical tests of 
paragraph (d) of this appendix, the two long edges and the upper edge 
are held securely; (2) in the horizontal test of paragraph (e) of this 
appendix, the two long edges and the edge away from the flame are held 
securely; (3) the exposed area of the specimen is at least 2 inches wide 
and 12 inches long, unless the actual size used in the airplane is 
smaller; and (4) the edge to which the burner flame is applied must not 
consist of the finished or protected edge of the specimen but must be 
representative of the actual cross section of the material or part 
installed in the airplane. When performing the test prescribed in 
paragraph (f)

[[Page 345]]

of this appendix, the specimen must be mounted in metal frame so that 
all four edges are held securely and the exposed area of the specimen is 
at least 8 inches by 8 inches.
    (c) Apparatus. Except as provided in paragraph (g) of this appendix, 
tests must be conducted in a draft-free cabinet in accordance with 
Federal Test Method Standard 191 Method 5903 (revised Method 5902) which 
is available from the General Services Administration, Business Service 
Center, Region 3, Seventh and D Streets SW., Washington, D.C. 20407, or 
with some other approved equivalent method. Specimens which are too 
large for the cabinet must be tested in similar draft-free conditions.
    (d) Vertical test. A minimum of three specimens must be tested and 
the results averaged. For fabrics, the direction of weave corresponding 
to the most critical flammability conditions must be parallel to the 
longest dimension. Each specimen must be supported vertically. The 
specimen must be exposed to a Bunsen or Tirrill burner with a nominal 
\3/8\-inch I.D. tube adjusted to give a flame of 1\1/2\ inches in 
height. The minimum flame temperature measured by a calibrated 
thermocouple pryometer in the center of the flame must be 1550 [deg]F. 
The lower edge of the specimen must be three-fourths inch above the top 
edge of the burner. The flame must be applied to the center line of the 
lower edge of the specimen. For materials covered by Sec. Sec. 
23.853(d)(3)(i) and 23.853(f), the flame must be applied for 60 seconds 
and then removed. For materials covered by Sec. 23.853(d)(3)(ii), the 
flame must be applied for 12 seconds and then removed. Flame time, burn 
length, and flaming time of drippings, if any, must be recorded. The 
burn length determined in accordance with paragraph (h) of this appendix 
must be measured to the nearest one-tenth inch.
    (e) Horizontal test. A minimum of three specimens must be tested and 
the results averaged. Each specimen must be supported horizontally. The 
exposed surface when installed in the airplane must be face down for the 
test. The specimen must be exposed to a Bunsen burner or Tirrill burner 
with a nominal \3/8\-inch I.D. tube adjusted to give a flame of 1\1/2\ 
inches in height. The minimum flame temperature measured by a calibrated 
thermocouple pyrometer in the center of the flame must be 1550 [deg]F. 
The specimen must be positioned so that the edge being tested is three-
fourths of an inch above the top of, and on the center line of, the 
burner. The flame must be applied for 15 seconds and then removed. A 
minimum of 10 inches of the specimen must be used for timing purposes, 
approximately 1\1/2\ inches must burn before the burning front reaches 
the timing zone, and the average burn rate must be recorded.
    (f) Forty-five degree test. A minimum of three specimens must be 
tested and the results averaged. The specimens must be supported at an 
angle of 45 degrees to a horizontal surface. The exposed surface when 
installed in the aircraft must be face down for the test. The specimens 
must be exposed to a Bunsen or Tirrill burner with a nominal \3/8\ inch 
I.D. tube adjusted to give a flame of 1\1/2\ inches in height. The 
minimum flame temperature measured by a calibrated thermocouple 
pyrometer in the center of the flame must be 1550 [deg]F. Suitable 
precautions must be taken to avoid drafts. The flame must be applied for 
30 seconds with one-third contacting the material at the center of the 
specimen and then removed. Flame time, glow time, and whether the flame 
penetrates (passes through) the specimen must be recorded.
    (g) Sixty-degree test. A minimum of three specimens of each wire 
specification (make and size) must be tested. The specimen of wire or 
cable (including insulation) must be placed at an angle of 60 degrees 
with the horizontal in the cabinet specified in paragraph (c) of this 
appendix, with the cabinet door open during the test or placed within a 
chamber approximately 2 feet high x 1 foot x 1 foot, open at the top and 
at one vertical side (front), that allows sufficient flow of air for 
complete combustion but is free from drafts. The specimen must be 
parallel to and approximately 6 inches from the front of the chamber. 
The lower end of the specimen must be held rigidly clamped. The upper 
end of the specimen must pass over a pulley or rod and must have an 
appropriate weight attached to it so that the specimen is held tautly 
throughout the flammability test. The test specimen span between lower 
clamp and upper pulley or rod must be 24 inches and must be marked 8 
inches from the lower end to indicate the central point for flame 
application. A flame from a Bunsen or Tirrill burner must be applied for 
30 seconds at the test mark. The burner must be mounted underneath the 
test mark on the specimen, perpendicular to the specimen and at an angle 
of 30 degrees to the vertical plane of the specimen. The burner must 
have a nominal bore of three-eighths inch, and must be adjusted to 
provide a three-inch-high flame with an inner cone approximately one-
third of the flame height. The minimum temperature of the hottest 
portion of the flame, as measured with a calibrated thermocouple 
pyrometer, may not be less than 1,750 [deg]F. The burner must be 
positioned so that the hottest portion of the flame is applied to the 
test mark on the wire. Flame time, burn length, and flaming time 
drippings, if any, must be recorded. The burn length determined in 
accordance with paragraph (h) of this appendix must be measured to the 
nearest one-tenth inch. Breaking of the wire specimen is not considered 
a failure.

[[Page 346]]

    (h) Burn length. Burn length is the distance from the original edge 
to the farthest evidence of damage to the test specimen due to flame 
impingement, including areas of partial or complete consumption, 
charring, or embrittlement, but not including areas sooted, stained, 
warped, or discolored, nor areas where material has shrunk or melted 
away from the heat source.

Part II--Test Method To Determine the Flammability and Flame Propagation 
        Characteristics of Thermal/Acoustic Insulation Materials

    Use this test method to evaluate the flammability and flame 
propagation characteristics of thermal/acoustic insulation when exposed 
to both a radiant heat source and a flame.
    (a) Definitions.
    Flame propagation means the furthest distance of the propagation of 
visible flame towards the far end of the test specimen, measured from 
the midpoint of the ignition source flame. Measure this distance after 
initially applying the ignition source and before all flame on the test 
specimen is extinguished. The measurement is not a determination of burn 
length made after the test.
    Radiant heat source means an electric or air propane panel.
    Thermal/acoustic insulation means a material or system of materials 
used to provide thermal and/or acoustic protection. Examples include 
fiberglass or other batting material encapsulated by a film covering and 
foams.
    Zero point means the point of application of the pilot burner to the 
test specimen.
    (b) Test apparatus.
    [GRAPHIC] [TIFF OMITTED] TR02DE11.087
    
    (1) Radiant panel test chamber. Conduct tests in a radiant panel 
test chamber (see figure F1 above). Place the test chamber under an 
exhaust hood to facilitate clearing the chamber of smoke after each 
test. The radiant panel test chamber must be an enclosure 55 inches 
(1397 mm) long by 19.5 inches (495 mm) deep by 28 inches (710 mm) to 30 
inches (maximum) (762 mm) above the test specimen. Insulate the sides, 
ends, and top with a fibrous ceramic insulation, such as Kaowool MTM 
board. On the front side, provide a 52 by 12-inch (1321 by 305 mm) 
draft-free, high-temperature, glass window for viewing the sample during 
testing. Place a door below the window to provide access to the movable 
specimen platform holder. The bottom of the test chamber must be a 
sliding steel platform that has provision for securing the test specimen 
holder in a fixed and level position. The chamber must have an internal 
chimney with exterior dimensions of 5.1 inches (129 mm) wide, by 16.2 
inches (411 mm) deep by 13 inches (330 mm) high at the opposite end of 
the chamber from the radiant energy source. The interior dimensions must 
be 4.5 inches (114 mm) wide by 15.6 inches (395 mm) deep. The chimney 
must extend to the top of the chamber (see figure F2).

[[Page 347]]

[GRAPHIC] [TIFF OMITTED] TR02DE11.088

    (2) Radiant heat source. Mount the radiant heat energy source in a 
cast iron frame or equivalent. An electric panel must have six, 3-inch 
wide emitter strips. The emitter strips must be perpendicular to the 
length of the panel. The panel must have a radiation surface of 12\7/8\ 
by 18\1/2\ inches (327 by 470 mm). The panel must be capable of 
operating at temperatures up to 1300 [deg]F (704 [deg]C). An air propane 
panel must be made of a porous refractory material and have a radiation 
surface of 12 by 18 inches (305 by 457 mm). The panel must be capable of 
operating at temperatures up to 1,500 [deg]F (816 [deg]C). See figures 
F3a and F3b.

[[Page 348]]

[GRAPHIC] [TIFF OMITTED] TR02DE11.089

    (i) Electric radiant panel. The radiant panel must be 3-phase and 
operate at 208 volts. A single-phase, 240 volt panel is also acceptable. 
Use a solid-state power controller and microprocessor-based controller 
to set the electric panel operating parameters.
    (ii) Gas radiant panel. Use propane (liquid petroleum gas--2.1 UN 
1075) for the radiant

[[Page 349]]

panel fuel. The panel fuel system must consist of a venturi-type 
aspirator for mixing gas and air at approximately atmospheric pressure. 
Provide suitable instrumentation for monitoring and controlling the flow 
of fuel and air to the panel. Include an air flow gauge, an air flow 
regulator, and a gas pressure gauge.
    (iii) Radiant panel placement. Mount the panel in the chamber at 30 
degrees to the horizontal specimen plane, and 7\1/2\ inches above the 
zero point of the specimen.
    (3) Specimen holding system.
    (i) The sliding platform serves as the housing for test specimen 
placement. Brackets may be attached (via wing nuts) to the top lip of 
the platform in order to accommodate various thicknesses of test 
specimens. Place the test specimens on a sheet of Kaowool MTM board or 
1260 Standard Board (manufactured by Thermal Ceramics and available in 
Europe), or equivalent, either resting on the bottom lip of the sliding 
platform or on the base of the brackets. It may be necessary to use 
multiple sheets of material based on the thickness of the test specimen 
(to meet the sample height requirement). Typically, these non-
combustible sheets of material are available in \1/4\-inch (6 mm) 
thicknesses. See figure F4. A sliding platform that is deeper than the 
2-inch (50.8mm) platform shown in figure F4 is also acceptable as long 
as the sample height requirement is met.
[GRAPHIC] [TIFF OMITTED] TR02DE11.090

    (ii) Attach a \1/2\-inch (13 mm) piece of Kaowool MTM board or other 
high temperature material measuring 41\1/2\ by 8\1/4\ inches (1054 by 
210 mm) to the back of the platform. This board serves as a heat 
retainer and protects the test specimen from excessive preheating. The 
height of this board may not impede the sliding platform movement (in 
and out of the test chamber). If the platform has been fabricated such 
that the back side of the platform is high enough to prevent excess 
preheating of the specimen when the sliding platform is out, a retainer 
board is not necessary.
    (iii) Place the test specimen horizontally on the non-combustible 
board(s). Place a steel retaining/securing frame fabricated of mild 
steel, having a thickness of \1/8\-inch (3.2 mm) and overall dimensions 
of 23 by 13\1/8\ inches (584 by 333 mm) with a specimen opening of 19 by 
10\3/4\ inches (483 by 273 mm) over the test specimen. The front, back, 
and right portions of the top flange of the frame must rest on the top 
of the sliding platform, and the bottom flanges must pinch all 4 sides 
of the test specimen. The right bottom flange must be flush with the 
sliding platform. See figure F5.

[[Page 350]]

[GRAPHIC] [TIFF OMITTED] TR02DE11.091

    (4) Pilot Burner. The pilot burner used to ignite the specimen must 
be a BernzomaticTM commercial propane venturi torch with an axially 
symmetric burner tip and a propane supply tube with an orifice diameter 
of 0.006 inches (0.15 mm). The length of the burner tube must be 2\7/8\ 
inches (71 mm). The propane flow must be adjusted via gas pressure 
through an in-line regulator to produce a blue inner cone length of \3/
4\-inch (19 mm). A \3/4\-inch (19 mm) guide (such as a thin strip of 
metal) may be soldered to the top of the burner to aid in setting the 
flame height. The overall flame length must be approximately 5 inches 
long (127 mm). Provide a way to move the burner out of the ignition 
position so that the flame is horizontal and at least 2 inches (50 mm) 
above the specimen plane. See figure F6.

[[Page 351]]

[GRAPHIC] [TIFF OMITTED] TR02DE11.092

    (5) Thermocouples. Install a 24 American Wire Gauge (AWG) Type K 
(Chromel- Alumel) thermocouple in the test chamber for temperature 
monitoring. Insert it into the chamber through a small hole drilled 
through the back of the chamber. Place the thermocouple so that it 
extends 11 inches (279 mm) out from the back of the chamber wall, 11\1/
2\ inches (292 mm) from the right side of the chamber wall, and is 2 
inches (51 mm) below the radiant panel. The use of other thermocouples 
is optional.
    (6) Calorimeter. The calorimeter must be a one-inch cylindrical 
water-cooled, total heat flux density, foil type Gardon Gage that has a 
range of 0 to 5 BTU/ft \2\-second (0 to 5.7 Watts/cm \2\).
    (7) Calorimeter calibration specification and procedure.
    (i) Calorimeter specification.
    (A) Foil diameter must be 0.25 0.005 inches 
(6.35 0.13 mm).
    (B) Foil thickness must be 0.0005 0.0001 
inches (0.013 0.0025 mm).
    (C) Foil material must be thermocouple grade Constantan.
    (D) Temperature measurement must be a Copper Constantan 
thermocouple.
    (E) The copper center wire diameter must be 0.0005 inches (0.013 
mm).
    (F) The entire face of the calorimeter must be lightly coated with 
``Black Velvet'' paint having an emissivity of 96 or greater.
    (ii) Calorimeter calibration.
    (A) The calibration method must be by comparison to a like 
standardized transducer.
    (B) The standardized transducer must meet the specifications given 
in paragraph II(b)(6) of this appendix.
    (C) Calibrate the standard transducer against a primary standard 
traceable to the National Institute of Standards and Technology (NIST).
    (D) The method of transfer must be a heated graphite plate.
    (E) The graphite plate must be electrically heated, have a clear 
surface area on each side of the plate of at least 2 by 2 inches (51 by 
51 mm), and be \1/8\-inch \1/16\-inch thick (3.2 
1.6 mm).
    (F) Center the 2 transducers on opposite sides of the plates at 
equal distances from the plate.
    (G) The distance of the calorimeter to the plate must be no less 
than 0.0625 inches (1.6 mm), and no greater than 0.375 inches (9.5 mm).
    (H) The range used in calibration must be at least 0-3.5 BTUs/ft 
\2\-second (0-3.9 Watts/cm \2\) and no greater than 0-5.7 BTUs/ft \2\-
second (0-6.4 Watts/cm \2\).
    (I) The recording device used must record the 2 transducers 
simultaneously or at least within \1/10\ of each other.
    (8) Calorimeter fixture. With the sliding platform pulled out of the 
chamber, install the calorimeter holding frame and place a sheet of non-
combustible material in the bottom of the sliding platform adjacent to 
the holding frame. This will prevent heat losses during calibration. The 
frame must be 13\1/8\ inches (333 mm) deep (front to back) by 8 inches 
(203 mm) wide and must rest on the top of the sliding platform. It must 
be fabricated of \1/8\-inch (3.2 mm) flat stock steel and have an 
opening that accommodates a \1/2\-inch (12.7 mm) thick piece of 
refractory board, which is level with the top of the sliding platform. 
The board must have three 1-inch (25.4 mm) diameter holes drilled 
through the board for calorimeter insertion. The distance to the radiant 
panel surface from the centerline of the first hole (``zero'' position) 
must be 7\1/2\ \1/8\-inches (191 3 mm). The distance between the centerline of the first 
hole to the centerline of the second hole must be 2 inches (51 mm). It 
must also be the same distance from the centerline of the second hole to 
the centerline of the third hole. See figure F7. A calorimeter holding 
frame that differs in construction is acceptable as long as the height 
from the centerline of the first hole to the radiant panel and the 
distance between holes is the same as described in this paragraph.

[[Page 352]]

[GRAPHIC] [TIFF OMITTED] TR02DE11.093

    (9) Instrumentation. Provide a calibrated recording device with an 
appropriate range or a computerized data acquisition system to measure 
and record the outputs of the calorimeter and the thermocouple. The data 
acquisition system must be capable of recording the calorimeter output 
every second during calibration.
    (10) Timing device. Provide a stopwatch or other device, accurate to 
1 second/hour, to measure the time of application 
of the pilot burner flame.
    (c) Test specimens.
    (1) Specimen preparation. Prepare and test a minimum of three test 
specimens. If an oriented film cover material is used, prepare and test 
both the warp and fill directions.
    (2) Construction. Test specimens must include all materials used in 
construction of the insulation (including batting, film, scrim, tape, 
etc.). Cut a piece of core material such as foam or fiberglass, and cut 
a piece of film cover material (if used) large enough to cover the core 
material. Heat sealing is the preferred method of preparing fiberglass 
samples, since they can be made without compressing the fiberglass 
(``box sample''). Cover materials that are not heat sealable may be 
stapled, sewn, or taped as long as the cover material is sufficiently 
over-cut to be drawn down the sides without compressing the core 
material. The fastening means should be as continuous as possible along 
the length of the seams. The specimen thickness must be of the same 
thickness as installed in the airplane.
    (3) Specimen Dimensions. To facilitate proper placement of specimens 
in the sliding platform housing, cut non-rigid core materials, such as 
fiberglass, 12\1/2\ inches (318mm) wide by 23 inches (584mm) long. Cut 
rigid materials, such as foam, 11\1/2\ \1/4\ 
inches (292 mm 6mm) wide by 23 inches (584mm) long 
in order to fit properly in the sliding platform housing and provide a 
flat, exposed surface equal to the opening in the housing.
    (d) Specimen conditioning. Condition the test specimens at 70 5 [deg]F (21 2 [deg]C) and 55 
percent 10 percent relative humidity, for a 
minimum of 24 hours prior to testing.
    (e) Apparatus Calibration.
    (1) With the sliding platform out of the chamber, install the 
calorimeter holding frame. Push the platform back into the chamber and 
insert the calorimeter into the first hole (``zero'' position). See 
figure F7. Close the bottom door located below the sliding platform. The 
distance from the centerline of the calorimeter to the radiant panel 
surface at this point must be 7\1/2\ inches \1/8\ 
(191 mm 3). Before igniting the radiant panel, 
ensure that the calorimeter face is clean and that there is water 
running through the calorimeter.
    (2) Ignite the panel. Adjust the fuel/air mixture to achieve 1.5 
BTUs/feet\2\-second 5 percent (1.7 Watts/cm\2\ 
5 percent) at the ``zero'' position. If using an 
electric panel, set the power controller to achieve the proper heat 
flux. Allow the unit to reach steady state (this may take up to 1 hour). 
The pilot burner must be off and in the down position during this time.

[[Page 353]]

    (3) After steady-state conditions have been reached, move the 
calorimeter 2 inches (51 mm) from the ``zero'' position (first hole) to 
position 1 and record the heat flux. Move the calorimeter to position 2 
and record the heat flux. Allow enough time at each position for the 
calorimeter to stabilize. Table 1 depicts typical calibration values at 
the three positions.

                       Table 1--Calibration Table
------------------------------------------------------------------------
           Position               BTU/feet\2\ sec        Watts/cm\2\
------------------------------------------------------------------------
``Zero'' Position.............                 1.5                  1.7
Position 1....................      1.51-1.50-1.49       1.71-1.70-1.69
Position 2....................           1.43-1.44            1.62-1.63
------------------------------------------------------------------------

    (4) Open the bottom door, remove the calorimeter and holder fixture. 
Use caution as the fixture is very hot.
    (f) Test Procedure.
    (1) Ignite the pilot burner. Ensure that it is at least 2 inches (51 
mm) above the top of the platform. The burner may not contact the 
specimen until the test begins.
    (2) Place the test specimen in the sliding platform holder. Ensure 
that the test sample surface is level with the top of the platform. At 
``zero'' point, the specimen surface must be 7\1/2\ inches \1/8\ inch (191 mm 3) below the 
radiant panel.
    (3) Place the retaining/securing frame over the test specimen. It 
may be necessary (due to compression) to adjust the sample (up or down) 
in order to maintain the distance from the sample to the radiant panel 
(7\1/2\ inches \1/8\ inch (191 mm 3) at ``zero'' position). With film/fiberglass 
assemblies, it is critical to make a slit in the film cover to purge any 
air inside. This allows the operator to maintain the proper test 
specimen position (level with the top of the platform) and to allow 
ventilation of gases during testing. A longitudinal slit, approximately 
2 inches (51mm) in length, must be centered 3 inches \1/2\ inch (76mm 13mm) from the 
left flange of the securing frame. A utility knife is acceptable for 
slitting the film cover.
    (4) Immediately push the sliding platform into the chamber and close 
the bottom door.
    (5) Bring the pilot burner flame into contact with the center of the 
specimen at the ``zero'' point and simultaneously start the timer. The 
pilot burner must be at a 27 degree angle with the sample and be 
approximately \1/2\ inch (12 mm) above the sample. See figure F7. A 
stop, as shown in figure F8, allows the operator to position the burner 
correctly each time.
[GRAPHIC] [TIFF OMITTED] TR02DE11.094


[[Page 354]]


    (6) Leave the burner in position for 15 seconds and then remove to a 
position at least 2 inches (51 mm) above the specimen.
    (g) Report.
    (1) Identify and describe the test specimen.
    (2) Report any shrinkage or melting of the test specimen.
    (3) Report the flame propagation distance. If this distance is less 
than 2 inches, report this as a pass (no measurement required).
    (4) Report the after-flame time.
    (h) Requirements.
    (1) There must be no flame propagation beyond 2 inches (51 mm) to 
the left of the centerline of the pilot flame application.
    (2) The flame time after removal of the pilot burner may not exceed 
3 seconds on any specimen.

[Amdt. 23-23, 43 FR 50594, Oct. 30, 1978, as amended by Amdt. 23-34, 52 
FR 1835, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23-49, 61 FR 
5170, Feb. 9, 1996; Amdt. 23-62, 76 FR 75763, Dec. 2, 2011]



  Sec. Appendix G to Part 23--Instructions for Continued Airworthiness

    G23.1 General. (a) This appendix specifies requirements for the 
preparation of Instructions for Continued Airworthiness as required by 
Sec. 23.1529.
    (b) The Instructions for Continued Airworthiness for each airplane 
must include the Instructions for Continued Airworthiness for each 
engine and propeller (hereinafter designated `products'), for each 
appliance required by this chapter, and any required information 
relating to the interface of those appliances and products with the 
airplane. If Instructions for Continued Airworthiness are not supplied 
by the manufacturer of an appliance or product installed in the 
airplane, the Instructions for Continued Airworthiness for the airplane 
must include the information essential to the continued airworthiness of 
the airplane.
    (c) The applicant must submit to the FAA a program to show how 
changes to the Instructions for Continued Airworthiness made by the 
applicant or by the manufacturers of products and appliances installed 
in the airplane will be distributed.
    G23.2 Format. (a) The Instructions for Continued Airworthiness must 
be in the form of a manual or manuals as appropriate for the quantity of 
data to be provided.
    (b) The format of the manual or manuals must provide for a practical 
arrangement.
    G23.3 Content. The contents of the manual or manuals must be 
prepared in the English language. The Instructions for Continued 
Airworthiness must contain the following manuals or sections, as 
appropriate, and information:
    (a) Airplane maintenance manual or section. (1) Introduction 
information that includes an explanation of the airplane's features and 
data to the extent necessary for maintenance or preventive maintenance.
    (2) A description of the airplane and its systems and installations 
including its engines, propellers, and appliances.
    (3) Basic control and operation information describing how the 
airplane components and systems are controlled and how they operate, 
including any special procedures and limitations that apply.
    (4) Servicing information that covers details regarding servicing 
points, capacities of tanks, reservoirs, types of fluids to be used, 
pressures applicable to the various systems, location of access panels 
for inspection and servicing, locations of lubrication points, 
lubricants to be used, equipment required for servicing, tow 
instructions and limitations, mooring, jacking, and leveling 
information.
    (b) Maintenance instructions. (1) Scheduling information for each 
part of the airplane and its engines, auxiliary power units, propellers, 
accessories, instruments, and equipment that provides the recommended 
periods at which they should be cleaned, inspected, adjusted, tested, 
and lubricated, and the degree of inspection, the applicable wear 
tolerances, and work recommended at these periods. However, the 
applicant may refer to an accessory, instrument, or equipment 
manufacturer as the source of this information if the applicant shows 
that the item has an exceptionally high degree of complexity requiring 
specialized maintenance techniques, test equipment, or expertise. The 
recommended overhaul periods and necessary cross reference to the 
Airworthiness Limitations section of the manual must also be included. 
In addition, the applicant must include an inspection program that 
includes the frequency and extent of the inspections necessary to 
provide for the continued airworthiness of the airplane.
    (2) Troubleshooting information describing probable malfunctions, 
how to recognize those malfunctions, and the remedial action for those 
malfunctions.
    (3) Information describing the order and method of removing and 
replacing products and parts with any necessary precautions to be taken.
    (4) Other general procedural instructions including procedures for 
system testing during ground running, symmetry checks, weighing and 
determining the center of gravity, lifting and shoring, and storage 
limitations.
    (c) Diagrams of structural access plates and information needed to 
gain access for inspections when access plates are not provided.
    (d) Details for the application of special inspection techniques 
including radiographic and ultrasonic testing where such processes are 
specified.

[[Page 355]]

    (e) Information needed to apply protective treatments to the 
structure after inspection.
    (f) All data relative to structural fasteners such as 
identification, discard recommendations, and torque values.
    (g) A list of special tools needed.
    (h) In addition, for commuter category airplanes, the following 
information must be furnished:
    (1) Electrical loads applicable to the various systems;
    (2) Methods of balancing control surfaces;
    (3) Identification of primary and secondary structures; and
    (4) Special repair methods applicable to the airplane.
    G23.4 Airworthiness Limitations section. The Instructions for 
Continued Airworthiness must contain a section titled Airworthiness 
Limitations that is segregated and clearly distinguishable from the rest 
of the document. This section must set forth each mandatory replacement 
time, structural inspection interval, and related structural inspection 
procedure required for type certification. If the Instructions for 
Continued Airworthiness consist of multiple documents, the section 
required by this paragraph must be included in the principal manual. 
This section must contain a legible statement in a prominent location 
that reads: ``The Airworthiness Limitations section is FAA approved and 
specifies maintenance required under Sec. Sec. 43.16 and 91.403 of the 
Federal Aviation Regulations unless an alternative program has been FAA 
approved.''

[Amdt. 23-26, 45 FR 60171, Sept. 11, 1980, as amended by Amdt. 23-34, 52 
FR 1835, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23-37, 54 FR 
34329, Aug. 18, 1989]



 Sec. Appendix H to Part 23--Installation of An Automatic Power Reserve 
                              (APR) System

    H23.1, General.
    (a) This appendix specifies requirements for installation of an APR 
engine power control system that automatically advances power or thrust 
on the operating engine(s) in the event any engine fails during takeoff.
    (b) With the APR system and associated systems functioning normally, 
all applicable requirements (except as provided in this appendix) must 
be met without requiring any action by the crew to increase power or 
thrust.
    H23.2, Definitions.
    (a) Automatic power reserve system means the entire automatic system 
used only during takeoff, including all devices both mechanical and 
electrical that sense engine failure, transmit signals, actuate fuel 
controls or power levers on operating engines, including power sources, 
to achieve the scheduled power increase and furnish cockpit information 
on system operation.
    (b) Selected takeoff power, notwithstanding the definition of 
``Takeoff Power'' in part 1 of the Federal Aviation Regulations, means 
the power obtained from each initial power setting approved for takeoff.
    (c) Critical Time Interval, as illustrated in figure H1, means that 
period starting at V1 minus one second and ending at the 
intersection of the engine and APR failure flight path line with the 
minimum performance all engine flight path line. The engine and APR 
failure flight path line intersects the one-engine-inoperative flight 
path line at 400 feet above the takeoff surface. The engine and APR 
failure flight path is based on the airplane's performance and must have 
a positive gradient of at least 0.5 percent at 400 feet above the 
takeoff surface.

[[Page 356]]

[GRAPHIC] [TIFF OMITTED] TC28SE91.030

    H23.3, Reliability and performance requirements.
    (a) It must be shown that, during the critical time interval, an APR 
failure that increases or does not affect power on either engine will 
not create a hazard to the airplane, or it must be shown that such 
failures are improbable.
    (b) It must be shown that, during the critical time interval, there 
are no failure modes of the APR system that would result in a failure 
that will decrease the power on either engine or it must be shown that 
such failures are extremely improbable.
    (c) It must be shown that, during the critical time interval, there 
will be no failure of the APR system in combination with an engine 
failure or it must be shown that such failures are extremely improbable.
    (d) All applicable performance requirements must be met with an 
engine failure occurring at the most critical point during takeoff with 
the APR system functioning normally.
    H23.4, Power setting.
    The selected takeoff power set on each engine at the beginning of 
the takeoff roll may not be less than--
    (a) The power necessary to attain, at V1, 90 percent of 
the maximum takeoff power approved for the airplane for the existing 
conditions;
    (b) That required to permit normal operation of all safety-related 
systems and equipment that are dependent upon engine power or power 
lever position; and
    (c) That shown to be free of hazardous engine response 
characteristics when power is advanced from the selected takeoff power 
level to the maximum approved takeoff power.
    H23.5, Powerplant controls--general.
    (a) In addition to the requirements of Sec. 23.1141, no single 
failure or malfunction (or probable combination thereof) of the APR, 
including associated systems, may cause the failure of any powerplant 
function necessary for safety.
    (b) The APR must be designed to--
    (1) Provide a means to verify to the flight crew before takeoff that 
the APR is in an operating condition to perform its intended function;
    (2) Automatically advance power on the operating engines following 
an engine failure during takeoff to achieve the maximum attainable 
takeoff power without exceeding engine operating limits;
    (3) Prevent deactivation of the APR by manual adjustment of the 
power levers following an engine failure;
    (4) Provide a means for the flight crew to deactivate the automatic 
function. This means must be designed to prevent inadvertent 
deactivation; and
    (5) Allow normal manual decrease or increase in power up to the 
maximum takeoff power approved for the airplane under the existing 
conditions through the use of power levers, as stated in Sec. 
23.1141(c), except as provided under paragraph (c) of H23.5 of this 
appendix.
    (c) For airplanes equipped with limiters that automatically prevent 
engine operating limits from being exceeded, other means may be used to 
increase the maximum level of power controlled by the power levers in 
the event of an APR failure. The means must be located on or forward of 
the power levers, must be easily identified and operated under all 
operating conditions by a single action of any pilot with the hand that 
is normally

[[Page 357]]

used to actuate the power levers, and must meet the requirements of 
Sec. 23.777 (a), (b), and (c).
    H23.6, Powerplant instruments.
    In addition to the requirements of Sec. 23.1305:
    (a) A means must be provided to indicate when the APR is in the 
armed or ready condition.
    (b) If the inherent flight characteristics of the airplane do not 
provide warning that an engine has failed, a warning system independent 
of the APR must be provided to give the pilot a clear warning of any 
engine failure during takeoff.
    (c) Following an engine failure at V1 or above, there 
must be means for the crew to readily and quickly verify that the APR 
has operated satisfactorily.

[Doc. No. 26344, 58 FR 18979, Apr. 9, 1993]

[[Page 358]]

                  Appendix I to Part 23--Seaplane Loads
[GRAPHIC] [TIFF OMITTED] TC28SE91.031


[[Page 359]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.032


[Amdt. 23-45, 58 FR 42167, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]



 Sec. Appendix J to Part 23--HIRF Environments and Equipment HIRF Test 
                                 Levels

    This appendix specifies the HIRF environments and equipment HIRF 
test levels for electrical and electronic systems under Sec. 23.1308. 
The field strength values for the HIRF environments and equipment HIRF 
test levels are expressed in root-mean-square units measured during the 
peak of the modulation cycle.
    (a) HIRF environment I is specified in the following table:

[[Page 360]]



                      Table I.--HIRF Environment I
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-2 MHz......................................         50         50
2 MHz-30 MHz......................................        100        100
30 MHz-100 MHz....................................         50         50
100 MHz-400 MHz...................................        100        100
400 MHz-700 MHz...................................        700         50
700 MHz-1 GHz.....................................        700        100
GHz-2 GHz.........................................      2,000        200
2 GHz-6 GHz.......................................      3,000        200
6 GHz-8 GHz.......................................      1,000        200
8 GHz-12 GHz......................................      3,000        300
12 GHz-18 GHz.....................................      2,000        200
18 GHz-40 GHz.....................................        600        200
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (b) HIRF environment II is specified in the following table:

                      Table II.-HIRF Environment II
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-500 kHz....................................         20         20
500 kHz-2 MHz.....................................         30         30
2 MHz-30 MHz......................................        100        100
30 MHz-100 MHz....................................         10         10
100 MHz-200 MHz...................................         30         10
200 MHz-400 MHz...................................         10         10
400 MHz-1 GHz.....................................        700         40
1 GHz-2 GHz.......................................      1,300        160
2 GHz-4 GHz.......................................      3,000        120
4 GHz-6 GHz.......................................      3,000        160
6 GHz-8 GHz.......................................        400        170
8 GHz-12 GHz......................................      1,230        230
12 GHz-18 GHz.....................................        730        190
18 GHz-40 GHz.....................................        600        150
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (c) Equipment HIRF Test Level 1. (1) From 10 kilohertz (kHz) to 400 
megahertz (MHz), use conducted susceptibility tests with continuous wave 
(CW) and 1 kHz square wave modulation with 90 percent depth or greater. 
The conducted susceptibility current must start at a minimum of 0.6 
milliamperes (mA) at 10 kHz, increasing 20 decibels (dB) per frequency 
decade to a minimum of 30 mA at 500 kHz.
    (2) From 500 kHz to 40 MHz, the conducted susceptibility current 
must be at least 30 mA.
    (3) From 40 MHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 30 mA at 40 MHz, decreasing 20 dB per frequency 
decade to a minimum of 3 mA at 400 MHz.
    (4) From 100 MHz to 400 MHz, use radiated susceptibility tests at a 
minimum of 20 volts per meter (V/m) peak with CW and 1 kHz square wave 
modulation with 90 percent depth or greater.
    (5) From 400 MHz to 8 gigahertz (GHz), use radiated susceptibility 
tests at a minimum of 150 V/m peak with pulse modulation of 4 percent 
duty cycle with a 1 kHz pulse repetition frequency. This signal must be 
switched on and off at a rate of 1 Hz with a duty cycle of 50 percent.
    (d) Equipment HIRF Test Level 2. Equipment HIRF test level 2 is HIRF 
environment II in table II of this appendix reduced by acceptable 
aircraft transfer function and attenuation curves. Testing must cover 
the frequency band of 10 kHz to 8 GHz.
    (e) Equipment HIRF Test Level 3. (1) From 10 kHz to 400 MHz, use 
conducted susceptibility tests, starting at a minimum of 0.15 mA at 10 
kHz, increasing 20 dB per frequency decade to a minimum of 7.5 mA at 500 
kHz.
    (2) From 500 kHz to 40 MHz, use conducted susceptibility tests at a 
minimum of 7.5 mA.
    (3) From 40 MHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 7.5 mA at 40 MHz, decreasing 20 dB per 
frequency decade to a minimum of 0.75 mA at 400 MHz.
    (4) From 100 MHz to 8 GHz, use radiated susceptibility tests at a 
minimum of 5 V/m.

[Doc. No. FAA-2006-23657, 72 FR 44025, Aug. 6, 2007]



PART 25_AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES--
Table of Contents



Special Federal Aviation Regulation No. 13
Special Federal Aviation Regulation No. 109

                            Subpart A_General

Sec.
25.1 Applicability.
25.2 Special retroactive requirements.
25.3 Special provisions for ETOPS type design approvals.
25.5 Incorporations by reference.

                            Subpart B_Flight

                                 General

25.21 Proof of compliance.
25.23 Load distribution limits.
25.25 Weight limits.
25.27 Center of gravity limits.
25.29 Empty weight and corresponding center of gravity.
25.31 Removable ballast.
25.33 Propeller speed and pitch limits.

                               Performance

25.101 General.
25.103 Stall speed.
25.105 Takeoff.
25.107 Takeoff speeds.

[[Page 361]]

25.109 Accelerate-stop distance.
25.111 Takeoff path.
25.113 Takeoff distance and takeoff run.
25.115 Takeoff flight path.
25.117 Climb: general.
25.119 Landing climb: All-engines-operating.
25.121 Climb: One-engine-inoperative.
25.123 En route flight paths.
25.125 Landing.

                   Controllability and Maneuverability

25.143 General.
25.145 Longitudinal control.
25.147 Directional and lateral control.
25.149 Minimum control speed.

                                  Trim

25.161 Trim.

                                Stability

25.171 General.
25.173 Static longitudinal stability.
25.175 Demonstration of static longitudinal stability.
25.177 Static lateral-directional stability.
25.181 Dynamic stability.

                                 Stalls

25.201 Stall demonstration.
25.203 Stall characteristics.
25.207 Stall warning.

                Ground and Water Handling Characteristics

25.231 Longitudinal stability and control.
25.233 Directional stability and control.
25.235 Taxiing condition.
25.237 Wind velocities.
25.239 Spray characteristics, control, and stability on water.

                    Miscellaneous Flight Requirements

25.251 Vibration and buffeting.
25.253 High-speed characteristics.
25.255 Out-of-trim characteristics.

                           Subpart C_Structure

                                 General

25.301 Loads.
25.303 Factor of safety.
25.305 Strength and deformation.
25.307 Proof of structure.

                              Flight Loads

25.321 General.

                   Flight Maneuver and Gust Conditions

25.331 Symmetric maneuvering conditions.
25.333 Flight maneuvering envelope.
25.335 Design airspeeds.
25.337 Limit maneuvering load factors.
25.341 Gust and turbulence loads.
25.343 Design fuel and oil loads.
25.345 High lift devices.
25.349 Rolling conditions.
25.351 Yaw maneuver conditions.

                        Supplementary Conditions

25.361 Engine torque.
25.362 Engine failure loads.
25.363 Side load on engine and auxiliary power unit mounts.
25.365 Pressurized compartment loads.
25.367 Unsymmetrical loads due to engine failure.
25.371 Gyroscopic loads.
25.373 Speed control devices.

                    Control Surface and System Loads

25.391 Control surface loads: General.
25.393 Loads parallel to hinge line.
25.395 Control system.
25.397 Control system loads.
25.399 Dual control system.
25.405 Secondary control system.
25.407 Trim tab effects.
25.409 Tabs.
25.415 Ground gust conditions.
25.427 Unsymmetrical loads.
25.445 Auxiliary aerodynamic surfaces.
25.457 Wing flaps.
25.459 Special devices.

                              Ground Loads

25.471 General.
25.473 Landing load conditions and assumptions.
25.477 Landing gear arrangement.
25.479 Level landing conditions.
25.481 Tail-down landing conditions.
25.483 One-gear landing conditions.
25.485 Side load conditions.
25.487 Rebound landing condition.
25.489 Ground handling conditions.
25.491 Taxi, takeoff and landing roll.
25.493 Braked roll conditions.
25.495 Turning.
25.497 Tail-wheel yawing.
25.499 Nose-wheel yaw and steering.
25.503 Pivoting.
25.507 Reversed braking.
25.509 Towing loads.
25.511 Ground load: unsymmetrical loads on multiple-wheel units.
25.519 Jacking and tie-down provisions.

                               Water Loads

25.521 General.
25.523 Design weights and center of gravity positions.
25.525 Application of loads.
25.527 Hull and main float load factors.
25.529 Hull and main float landing conditions.
25.531 Hull and main float takeoff condition.
25.533 Hull and main float bottom pressures.
25.535 Auxiliary float loads.
25.537 Seawing loads.

[[Page 362]]

                      Emergency Landing Conditions

25.561 General.
25.562 Emergency landing dynamic conditions.
25.563 Structural ditching provisions.

                           Fatigue Evaluation

25.571 Damage--tolerance and fatigue evaluation of structure.

                          Lightning Protection

25.581 Lightning protection.

                    Subpart D_Design and Construction

                                 General

25.601 General.
25.603 Materials.
25.605 Fabrication methods.
25.607 Fasteners.
25.609 Protection of structure.
25.611 Accessibility provisions.
25.613 Material strength properties and material design values.
25.619 Special factors.
25.621 Casting factors.
25.623 Bearing factors.
25.625 Fitting factors.
25.629 Aeroelastic stability requirements.
25.631 Bird strike damage.

                            Control Surfaces

25.651 Proof of strength.
25.655 Installation.
25.657 Hinges.

                             Control Systems

25.671 General.
25.672 Stability augmentation and automatic and power-operated systems.
25.675 Stops.
25.677 Trim systems.
25.679 Control system gust locks.
25.681 Limit load static tests.
25.683 Operation tests.
25.685 Control system details.
25.689 Cable systems.
25.693 Joints.
25.697 Lift and drag devices, controls.
25.699 Lift and drag device indicator.
25.701 Flap and slat interconnection.
25.703 Takeoff warning system.

                              Landing Gear

25.721 General.
25.723 Shock absorption tests.
25.725-25.727 [Reserved]
25.729 Retracting mechanism.
25.731 Wheels.
25.733 Tires.
25.735 Brakes and braking systems.
25.737 Skis.

                            Floats and Hulls

25.751 Main float buoyancy.
25.753 Main float design.
25.755 Hulls.

                   Personnel and Cargo Accommodations

25.771 Pilot compartment.
25.772 Pilot compartment doors.
25.773 Pilot compartment view.
25.775 Windshields and windows.
25.777 Cockpit controls.
25.779 Motion and effect of cockpit controls.
25.781 Cockpit control knob shape.
25.783 Fuselage doors.
25.785 Seats, berths, safety belts, and harnesses.
25.787 Stowage compartments.
25.789 Retention of items of mass in passenger and crew compartments and 
          galleys.
25.791 Passenger information signs and placards.
25.793 Floor surfaces.
25.795 Security considerations.

                          Emergency Provisions

25.801 Ditching.
25.803 Emergency evacuation.
25.807 Emergency exits.
25.809 Emergency exit arrangement.
25.810 Emergency egress assist means and escape routes.
25.811 Emergency exit marking.
25.812 Emergency lighting.
25.813 Emergency exit access.
25.815 Width of aisle.
25.817 Maximum number of seats abreast.
25.819 Lower deck service compartments (including galleys).
25.820 Lavatory doors.

                         Ventilation and Heating

25.831 Ventilation.
25.832 Cabin ozone concentration.
25.833 Combustion heating systems.

                             Pressurization

25.841 Pressurized cabins.
25.843 Tests for pressurized cabins.

                             Fire Protection

25.851 Fire extinguishers.
25.853 Compartment interiors.
25.854 Lavatory fire protection.
25.855 Cargo or baggage compartments.
25.856 Thermal/Acoustic insulation materials.
25.857 Cargo compartment classification.
25.858 Cargo or baggage compartment smoke or fire detection systems.
25.859 Combustion heater fire protection.
25.863 Flammable fluid fire protection.
25.865 Fire protection of flight controls, engine mounts, and other 
          flight structure.
25.867 Fire protection: other components.
25.869 Fire protection: systems.

[[Page 363]]

                              Miscellaneous

25.871 Leveling means.
25.875 Reinforcement near propellers.
25.899 Electrical bonding and protection against static electricity.

                          Subpart E_Powerplant

                                 General

25.901 Installation.
25.903 Engines.
25.904 Automatic takeoff thrust control system (ATTCS).
25.905 Propellers.
25.907 Propeller vibration and fatigue.
25.925 Propeller clearance.
25.929 Propeller deicing.
25.933 Reversing systems.
25.934 Turbojet engine thrust reverser system tests.
25.937 Turbopropeller-drag limiting systems.
25.939 Turbine engine operating characteristics.
25.941 Inlet, engine, and exhaust compatibility.
25.943 Negative acceleration.
25.945 Thrust or power augmentation system.

                               Fuel System

25.951 General.
25.952 Fuel system analysis and test.
25.953 Fuel system independence.
25.954 Fuel system lightning protection.
25.955 Fuel flow.
25.957 Flow between interconnected tanks.
25.959 Unusable fuel supply.
25.961 Fuel system hot weather operation.
25.963 Fuel tanks: general.
25.965 Fuel tank tests.
25.967 Fuel tank installations.
25.969 Fuel tank expansion space.
25.971 Fuel tank sump.
25.973 Fuel tank filler connection.
25.975 Fuel tank vents and carburetor vapor vents.
25.977 Fuel tank outlet.
25.979 Pressure fueling system.
25.981 Fuel tank ignition prevention.

                         Fuel System Components

25.991 Fuel pumps.
25.993 Fuel system lines and fittings.
25.994 Fuel system components.
25.995 Fuel valves.
25.997 Fuel strainer or filter.
25.999 Fuel system drains.
25.1001 Fuel jettisoning system.

                               Oil System

25.1011 General.
25.1013 Oil tanks.
25.1015 Oil tank tests.
25.1017 Oil lines and fittings.
25.1019 Oil strainer or filter.
25.1021 Oil system drains.
25.1023 Oil radiators.
25.1025 Oil valves.
25.1027 Propeller feathering system.

                                 Cooling

25.1041 General.
25.1043 Cooling tests.
25.1045 Cooling test procedures.

                            Induction System

25.1091 Air induction.
25.1093 Induction system icing protection.
25.1101 Carburetor air preheater design.
25.1103 Induction system ducts and air duct systems.
25.1105 Induction system screens.
25.1107 Inter-coolers and after-coolers.

                             Exhaust System

25.1121 General.
25.1123 Exhaust piping.
25.1125 Exhaust heat exchangers.
25.1127 Exhaust driven turbo-superchargers.

                   Powerplant Controls and Accessories

25.1141 Powerplant controls: general.
25.1142 Auxiliary power unit controls.
25.1143 Engine controls.
25.1145 Ignition switches.
25.1147 Mixture controls.
25.1149 Propeller speed and pitch controls.
25.1153 Propeller feathering controls.
25.1155 Reverse thrust and propeller pitch settings below the flight 
          regime.
25.1157 Carburetor air temperature controls.
25.1159 Supercharger controls.
25.1161 Fuel jettisoning system controls.
25.1163 Powerplant accessories.
25.1165 Engine ignition systems.
25.1167 Accessory gearboxes.

                       Powerplant Fire Protection

25.1181 Designated fire zones; regions included.
25.1182 Nacelle areas behind firewalls, and engine pod attaching 
          structures containing flammable fluid lines.
25.1183 Flammable fluid-carrying components.
25.1185 Flammable fluids.
25.1187 Drainage and ventilation of fire zones.
25.1189 Shutoff means.
25.1191 Firewalls.
25.1192 Engine accessory section diaphragm.
25.1193 Cowling and nacelle skin.
25.1195 Fire extinguishing systems.
25.1197 Fire extinguishing agents.
25.1199 Extinguishing agent containers.
25.1201 Fire extinguishing system materials.
25.1203 Fire detector system.
25.1207 Compliance.

[[Page 364]]

                           Subpart F_Equipment

                                 General

25.1301 Function and installation.
25.1302 Installed systems and equipment for use by the flightcrew.
25.1303 Flight and navigation instruments.
25.1305 Powerplant instruments.
25.1307 Miscellaneous equipment.
25.1309 Equipment, systems, and installations.
25.1310 Power source capacity and distribution.
25.1316 Electrical and electronic system lightning protection.
25.1317 High-intensity Radiated Fields (HIRF) Protection.

                        Instruments: Installation

25.1321 Arrangement and visibility.
25.1322 Flightcrew alerting.
25.1323 Airspeed indicating system.
25.1324 Angle of attack system.
25.1325 Static pressure systems.
25.1326 Pitot heat indication systems.
25.1327 Magnetic direction indicator.
25.1329 Flight guidance system.
25.1331 Instruments using a power supply.
25.1333 Instrument systems.
25.1337 Powerplant instruments.

                    Electrical Systems and Equipment

25.1351 General.
25.1353 Electrical equipment and installations.
25.1355 Distribution system.
25.1357 Circuit protective devices.
25.1360 Precautions against injury.
25.1362 Electrical supplies for emergency conditions.
25.1363 Electrical system tests.
25.1365 Electrical appliances, motors, and transformers.

                                 Lights

25.1381 Instrument lights.
25.1383 Landing lights.
25.1385 Position light system installation.
25.1387 Position light system dihedral angles.
25.1389 Position light distribution and intensities.
25.1391 Minimum intensities in the horizontal plane of forward and rear 
          position lights.
25.1393 Minimum intensities in any vertical plane of forward and rear 
          position lights.
25.1395 Maximum intensities in overlapping beams of forward and rear 
          position lights.
25.1397 Color specifications.
25.1399 Riding light.
25.1401 Anticollision light system.
25.1403 Wing icing detection lights.

                            Safety Equipment

25.1411 General.
25.1415 Ditching equipment.
25.1419 Ice protection.
25.1420 Supercooled large drop icing conditions.
25.1421 Megaphones.
25.1423 Public address system.

                         Miscellaneous Equipment

25.1431 Electronic equipment.
25.1433 Vacuum systems.
25.1435 Hydraulic systems.
25.1438 Pressurization and pneumatic systems.
25.1439 Protective breathing equipment.
25.1441 Oxygen equipment and supply.
25.1443 Minimum mass flow of supplemental oxygen.
25.1445 Equipment standards for the oxygen distributing system.
25.1447 Equipment standards for oxygen dispensing units.
25.1449 Means for determining use of oxygen.
25.1450 Chemical oxygen generators.
25.1453 Protection of oxygen equipment from rupture.
25.1455 Draining of fluids subject to freezing.
25.1457 Cockpit voice recorders.
25.1459 Flight data recorders.
25.1461 Equipment containing high energy rotors.

             Subpart G_Operating Limitations and Information

25.1501 General.

                          Operating Limitations

25.1503 Airspeed limitations: general.
25.1505 Maximum operating limit speed.
25.1507 Maneuvering speed.
25.1511 Flap extended speed.
25.1513 Minimum control speed.
25.1515 Landing gear speeds.
25.1516 Other speed limitations.
25.1517 Rough air speed, VRA.
25.1519 Weight, center of gravity, and weight distribution.
25.1521 Powerplant limitations.
25.1522 Auxiliary power unit limitations.
25.1523 Minimum flight crew.
25.1525 Kinds of operation.
25.1527 Ambient air temperature and operating altitude.
25.1529 Instructions for Continued Airworthiness.
25.1531 Maneuvering flight load factors.
25.1533 Additional operating limitations.
25.1535 ETOPS approval.

                          Markings and Placards

25.1541 General.
25.1543 Instrument markings: general.

[[Page 365]]

25.1545 Airspeed limitation information.
25.1547 Magnetic direction indicator.
25.1549 Powerplant and auxiliary power unit instruments.
25.1551 Oil quantity indication.
25.1553 Fuel quantity indicator.
25.1555 Control markings.
25.1557 Miscellaneous markings and placards.
25.1561 Safety equipment.
25.1563 Airspeed placard.

                         Airplane Flight Manual

25.1581 General.
25.1583 Operating limitations.
25.1585 Operating procedures.
25.1587 Performance information.

       Subpart H_Electrical Wiring Interconnection Systems (EWIS)

25.1701 Definition.
25.1703 Function and installation: EWIS.
25.1705 Systems and functions: EWIS.
25.1707 System separation: EWIS.
25.1709 System safety: EWIS.
25.1711 Component identification: EWIS.
25.1713 Fire protection: EWIS.
25.1715 Electrical bonding and protection against static electricity: 
          EWIS.
25.1717 Circuit protective devices: EWIS.
25.1719 Accessibility provisions: EWIS.
25.1721 Protection of EWIS.
25.1723 Flammable fluid fire protection: EWIS.
25.1725 Powerplants: EWIS.
25.1727 Flammable fluid shutoff means: EWIS.
25.1729 Instructions for Continued Airworthiness: EWIS.
25.1731 Powerplant and APU fire detector system: EWIS.
25.1733 Fire detector systems, general: EWIS.

             Subpart I_Special Federal Aviation Regulations

25.1801 SFAR No. 111--Lavatory Oxygen Systems.

Appendix A to Part 25
Appendix B to Part 25
Appendix C to Part 25
Appendix D to Part 25
Appendix E to Part 25
Appendix F to Part 25
Appendix G to Part 25--Continuous Gust Design Criteria
Appendix H to Part 25--Instructions for Continued Airworthiness
Appendix I to Part 25--Installation of an Automatic Takeoff Thrust 
          Control System (ATTCS)
Appendix J to Part 25--Emergency Evacuation
Appendix K to Part 25--Extended Operations (ETOPS)
Appendix L to Part 25--HIRF Environments and Equipment HIRF Test Levels
Appendix M to Part 25--Fuel Tank System Flammability Reduction Means
Appendix N to Part 25--Fuel Tank Flammability Exposure and Reliability 
          Analysis
Appendix O to Part 25--Supercooled Large Drop Icing Conditions

    Authority: 49 U.S.C. 106(g), 40113, 44701, 44702 and 44704.

    Source: Docket No. 5066, 29 FR 18291, Dec. 24, 1964, unless 
otherwise noted.



             Sec. Special Federal Aviation Regulation No. 13

    1. Applicability. Contrary provisions of the Civil Air Regulations 
regarding certification notwithstanding, \1\ this regulation shall 
provide the basis for approval by the Administrator of modifications of 
individual Douglas DC-3 and Lockheed L-18 airplanes subsequent to the 
effective date of this regulation.
---------------------------------------------------------------------------

    \1\ It is not intended to waive compliance with such airworthiness 
requirements as are included in the operating parts of the Civil Air 
Regulations for specific types of operation.
---------------------------------------------------------------------------

    2. General modifications. Except as modified in sections 3 and 4 of 
this regulation, an applicant for approval of modifications to a DC-3 or 
L-18 airplane which result in changes in design or in changes to 
approved limitations shall show that the modifications were accomplished 
in accordance with the rules of either Part 4a or Part 4b in effect on 
September 1, 1953, which are applicable to the modification being made: 
Provided, That an applicant may elect to accomplish a modification in 
accordance with the rules of Part 4b in effect on the date of 
application for the modification in lieu of Part 4a or Part 4b as in 
effect on September 1, 1953: And provided further, That each specific 
modification must be accomplished in accordance with all of the 
provisions contained in the elected rules relating to the particular 
modification.
    3. Specific conditions for approval. An applicant for any approval 
of the following specific changes shall comply with section 2 of this 
regulation as modified by the applicable provisions of this section.
    (a) Increase in take-off power limitation--1,200 to 1,350 
horsepower. The engine take-off power limitation for the airplane may be 
increased to more than 1,200 horsepower but not to more than 1,350 
horsepower per engine if the increase in power does not adversely affect 
the flight characteristics of the airplane.
    (b) Increase in take-off power limitation to more than 1,350 
horsepower. The engine take-

[[Page 366]]

off power limitation for the airplane may be increased to more than 
1,350 horsepower per engine if compliance is shown with the flight 
characteristics and ground handling requirements of Part 4b.
    (c) Installation of engines of not more than 1,830 cubic inches 
displacement and not having a certificated take-off rating of more than 
1,350 horsepower. Engines of not more than 1,830 cubic inches 
displacement and not having a certificated take-off rating of more than 
1,350 horsepower which necessitate a major modification of redesign of 
the engine installation may be installed, if the engine fire prevention 
and fire protection are equivalent to that on the prior engine 
installation.
    (d) Installation of engines of more than 1,830 cubic inches 
displacement or having certificated take-off rating of more than 1,350 
horsepower. Engines of more than 1,830 cubic inches displacement or 
having certificated take-off rating of more than 1,350 horsepower may be 
installed if compliance is shown with the engine installation 
requirements of Part 4b: Provided, That where literal compliance with 
the engine installation requirements of Part 4b is extremely difficult 
to accomplish and would not contribute materially to the objective 
sought, and the Administrator finds that the experience with the DC-3 or 
L-18 airplanes justifies it, he is authorized to accept such measures of 
compliance as he finds will effectively accomplish the basic objective.
    4. Establishment of new maximum certificated weights. An applicant 
for approval of new maximum certificated weights shall apply for an 
amendment of the airworthiness certificate of the airplane and shall 
show that the weights sought have been established, and the appropriate 
manual material obtained, as provided in this section.

    Note: Transport category performance requirements result in the 
establishment of maximum certificated weights for various altitudes.

    (a) Weights-25,200 to 26,900 for the DC-3 and 18,500 to 19,500 for 
the L-18. New maximum certificated weights of more than 25,200 but not 
more than 26,900 pounds for DC-3 and more than 18,500 but not more than 
19,500 pounds for L-18 airplanes may be established in accordance with 
the transport category performance requirements of either Part 4a or 
Part 4b, if the airplane at the new maximum weights can meet the 
structural requirements of the elected part.
    (b) Weights of more than 26,900 for the DC-3 and 19,500 for the L-
18. New maximum certificated weights of more than 26,900 pounds for DC-3 
and 19,500 pounds for L-18 airplanes shall be established in accordance 
with the structural performance, flight characteristics, and ground 
handling requirements of Part 4b: Provided, That where literal 
compliance with the structural requirements of Part 4b is extremely 
difficult to accomplish and would not contribute materially to the 
objective sought, and the Administrator finds that the experience with 
the DC-3 or L-18 airplanes justifies it, he is authorized to accept such 
measures of compliance as he finds will effectively accomplish the basic 
objective.
    (c) Airplane flight manual-performance operating information. An 
approved airplane flight manual shall be provided for each DC-3 and L-18 
airplane which has had new maximum certificated weights established 
under this section. The airplane flight manual shall contain the 
applicable performance information prescribed in that part of the 
regulations under which the new certificated weights were established 
and such additional information as may be necessary to enable the 
application of the take-off, en route, and landing limitations 
prescribed for transport category airplanes in the operating parts of 
the Civil Air Regulations.
    (d) Performance operating limitations. Each airplane for which new 
maximum certificated weights are established in accordance with 
paragraphs (a) or (b) of this section shall be considered a transport 
category airplane for the purpose of complying with the performance 
operating limitations applicable to the operations in which it is 
utilized.
    5. Reference. Unless otherwise provided, all references in this 
regulation to Part 4a and Part 4b are those parts of the Civil Air 
Regulations in effect on September 1, 1953.
    This regulation supersedes Special Civil Air Regulation SR-398 and 
shall remain effective until superseded or rescinded by the Board.

[19 FR 5039, Aug. 11, 1954. Redesignated at 29 FR 19099, Dec. 30, 1964]



            Sec. Special Federal Aviation Regulation No. 109

    1. Applicability. Contrary provisions of 14 CFR parts 21, 25, and 
119 of this chapter notwithstanding, an applicant is entitled to an 
amended type certificate or supplemental type certificate in the 
transport category, if the applicant complies with all applicable 
provisions of this SFAR.

                               Operations

    2. General.
    (a) The passenger capacity may not exceed 60. If more than 60 
passenger seats are installed, then:
    (1) If the extra seats are not suitable for occupancy during taxi, 
takeoff and landing, each extra seat must be clearly marked (e.g., a 
placard on the top of an armrest, or a placard sewn into the top of the 
back cushion) that the seat is not to be occupied during taxi, takeoff 
and landing.

[[Page 367]]

    (2) If the extra seats are suitable for occupancy during taxi, 
takeoff and landing (i.e., meet all the strength and passenger injury 
criteria in part 25), then a note must be included in the Limitations 
Section of the Airplane Flight Manual that there are extra seats 
installed but that the number of passengers on the airplane must not 
exceed 60. Additionally, there must be a placard installed adjacent to 
each door that can be used as a passenger boarding door that states that 
the maximum passenger capacity is 60. The placard must be clearly 
legible to passengers entering the airplane.
    (b) For airplanes outfitted with interior doors under paragraph 10 
of this SFAR, the airplane flight manual (AFM) must include an 
appropriate limitation that the airplane must be staffed with at least 
the following number of flight attendants who meet the requirements of 
14 CFR 91.533(b):
    (1) The number of flight attendants required by Sec. 91.533(a)(1) 
and (2) of this chapter, and
    (2) At least one flight attendant if the airplane model was 
originally certified for 75 passengers or more.
    (c) The AFM must include appropriate limitation(s) to require a 
preflight passenger briefing describing the appropriate functions to be 
performed by the passengers and the relevant features of the airplane to 
ensure the safety of the passengers and crew.
    (d) The airplane may not be offered for common carriage or operated 
for hire. The operating limitations section of the AFM must be revised 
to prohibit any operations involving the carriage of persons or property 
for compensation or hire. The operators may receive remuneration to the 
extent consistent with parts 125 and 91, subpart F, of this chapter.
    (e) A placard stating that ``Operations involving the carriage of 
persons or property for compensation or hire are prohibited,'' must be 
located in the area of the Airworthiness Certificate holder at the 
entrance to the flightdeck.
    (f) For passenger capacities of 45 to 60 passengers, analysis must 
be submitted that demonstrates that the airplane can be evacuated in 
less than 90 seconds under the conditions specified in Sec. 25.803 and 
appendix J to part 25.
    (g) In order for any airplane certified under this SFAR to be placed 
in part 135 or part 121 operations, the airplane must be brought back 
into full compliance with the applicable operational part.

                          Equipment and Design

    3. General. Unless otherwise noted, compliance is required with the 
applicable certification basis for the airplane. Some provisions of this 
SFAR impose alternative requirements to certain airworthiness standards 
that do not apply to airplanes certificated to earlier standards. Those 
airplanes with an earlier certification basis are not required to comply 
with those alternative requirements.
    4. Occupant Protection.
    (a) Firm Handhold. In lieu of the requirements of Sec. 25.785(j), 
there must be means provided to enable persons to steady themselves in 
moderately rough air while occupying aisles that are along the cabin 
sidewall, or where practicable, bordered by seats (seat backs providing 
a 25-pound minimum breakaway force are an acceptable means of 
compliance).
    (b) Injury criteria for multiple occupancy side-facing seats. The 
following requirements are only applicable to airplanes that are subject 
to Sec. 25.562.
    (1) Existing Criteria. All injury protection criteria of Sec. 
25.562(c)(1) through (c)(6) apply to the occupants of side-facing 
seating. The Head Injury Criterion (HIC) assessments are only required 
for head contact with the seat and/or adjacent structures.
    (2) Body-to-Body Contact. Contact between the head, pelvis, torso or 
shoulder area of one Anthropomorphic Test Dummy (ATD) with the head, 
pelvis, torso or shoulder area of the ATD in the adjacent seat is not 
allowed during the tests conducted in accordance with Sec. 25.562(b)(1) 
and (b)(2). Contact during rebound is allowed.
    (3) Thoracic Trauma. If the torso of an ATD at the forward-most seat 
place impacts the seat and/or adjacent structure during testing, 
compliance with the Thoracic Trauma Index (TTI) injury criterion must be 
substantiated by dynamic test or by rational analysis based on previous 
test(s) of a similar seat installation. TTI data must be acquired with a 
Side Impact Dummy (SID), as defined by 49 CFR part 572, subpart F, or an 
equivalent ATD or a more appropriate ATD and must be processed as 
defined in Federal Motor Vehicle Safety Standards (FMVSS) part 571.214, 
section S6.13.5 (49 CFR 571.214). The TTI must be less than 85, as 
defined in 49 CFR part 572, subpart F. Torso contact during rebound is 
acceptable and need not be measured.
    (4) Pelvis. If the pelvis of an ATD at any seat place impacts seat 
and/or adjacent structure during testing, pelvic lateral acceleration 
injury criteria must be substantiated by dynamic test or by rational 
analysis based on previous test(s) of a similar seat installation. 
Pelvic lateral acceleration may not exceed 130g. Pelvic acceleration 
data must be processed as defined in FMVSS part 571.214, section S6.13.5 
(49 CFR 571.214).
    (5) Body-to-Wall/Furnishing Contact. If the seat is installed aft of 
a structure--such as an interior wall or furnishing that may contact the 
pelvis, upper arm, chest, or head of an occupant seated next to the 
structure--the structure or a conservative representation of the 
structure and its stiffness must

[[Page 368]]

be included in the tests. It is recommended, but not required, that the 
contact surface of the actual structure be covered with at least two 
inches of energy absorbing protective padding (foam or equivalent) such 
as Ensolite.
    (6) Shoulder Strap Loads. Where upper torso straps (shoulder straps) 
are used for sofa occupants, the tension loads in individual straps may 
not exceed 1,750 pounds. If dual straps are used for restraining the 
upper torso, the total strap tension loads may not exceed 2,000 pounds.
    (7) Occupant Retention. All side-facing seats require end closures 
or other means to prevent the ATD's pelvis from translating beyond the 
end of the seat at any time during testing.
    (8) Test Parameters.
    (i) All seat positions need to be occupied by ATDs for the 
longitudinal tests.
    (ii) A minimum of one longitudinal test, conducted in accordance 
with the conditions specified in Sec. 25.562(b)(2), is required to 
assess the injury criteria as follows. Note that if a seat is installed 
aft of structure (such as an interior wall or furnishing) that does not 
have a homogeneous surface, an additional test or tests may be required 
to demonstrate that the injury criteria are met for the area which an 
occupant could contact. For example, different yaw angles could result 
in different injury considerations and may require separate tests to 
evaluate.
    (A) For configurations without structure (such as a wall or 
bulkhead) installed directly forward of the forward seat place, Hybrid 
II ATDs or equivalent must be in all seat places.
    (B) For configurations with structure (such as a wall or bulkhead) 
installed directly forward of the forward seat place, a side impact 
dummy or equivalent ATD or more appropriate ATD must be in the forward 
seat place and a Hybrid II ATD or equivalent must be in all other seat 
places.
    (C) The test may be conducted with or without deformed floor.
    (D) The test must be conducted with either no yaw or 10 degrees yaw 
for evaluating occupant injury. Deviating from the no yaw condition may 
not result in the critical area of contact not being evaluated. The 
upper torso restraint straps, where installed, must remain on the 
occupant's shoulder during the impact condition of Sec. 25.562(b)(2).
    (c) For the vertical test, conducted in accordance with the 
conditions specified in Sec. 25.562(b)(1), Hybrid II ATDs or equivalent 
must be used in all seat positions.
    5. Direct View. In lieu of the requirements of Sec. 25.785(h)(2), 
to the extent practical without compromising proximity to a required 
floor level emergency exit, the majority of installed flight attendant 
seats must be located to face the cabin area for which the flight 
attendant is responsible.
    6. Passenger Information Signs. Compliance with Sec. 25.791 is 
required except that for Sec. 25.791(a), when smoking is to be 
prohibited, notification to the passengers may be provided by a single 
placard so stating, to be conspicuously located inside the passenger 
compartment, easily visible to all persons entering the cabin in the 
immediate vicinity of each passenger entry door.
    7. Distance Between Exits. For an airplane that is required to 
comply with Sec. 25.807(f)(4), in effect as of July 24, 1989, which has 
more than one passenger emergency exit on each side of the fuselage, no 
passenger emergency exit may be more than 60 feet from any adjacent 
passenger emergency exit on the same side of the same deck of the 
fuselage, as measured parallel to the airplane's longitudinal axis 
between the nearest exit edges, unless the following conditions are met:
    (a) Each passenger seat must be located within 30 feet from the 
nearest exit on each side of the fuselage, as measured parallel to the 
airplane's longitudinal axis, between the nearest exit edge and the 
front of the seat bottom cushion.
    (b) The number of passenger seats located between two adjacent pairs 
of emergency exits (commonly referred to as a passenger zone) or between 
a pair of exits and a bulkhead or a compartment door (commonly referred 
to as a ``dead-end zone''), may not exceed the following:
    (1) For zones between two pairs of exits, 50 percent of the combined 
rated capacity of the two pairs of emergency exits.
    (2) For zones between one pair of exits and a bulkhead, 40 percent 
of the rated capacity of the pair of emergency exits.
    (c) The total number of passenger seats in the airplane may not 
exceed 33 percent of the maximum seating capacity for the airplane model 
using the exit ratings listed in Sec. 25.807(g) for the original 
certified exits or the maximum allowable after modification when exits 
are deactivated, whichever is less.
    (d) A distance of more than 60 feet between adjacent passenger 
emergency exits on the same side of the same deck of the fuselage, as 
measured parallel to the airplane's longitudinal axis between the 
nearest exit edges, is allowed only once on each side of the fuselage.
    8. Emergency Exit Signs. In lieu of the requirements of Sec. 
25.811(d)(1) and (2) a single sign at each exit may be installed 
provided:
    (a) The sign can be read from the aisle while directly facing the 
exit, and
    (b) The sign can be read from the aisle adjacent to the passenger 
seat that is farthest from the exit and that does not have an 
intervening bulkhead/divider or exit.
    9. Emergency Lighting.
    (a) Exit Signs. In lieu of the requirements of Sec. 25.812(b)(1), 
for airplanes that have a passenger seating configuration, excluding 
pilot seats, of 19 seats or less, the emergency exit

[[Page 369]]

signs required by Sec. 25.811(d)(1), (2), and (3) must have red letters 
at least 1-inch high on a white background at least 2 inches high. These 
signs may be internally electrically illuminated, or self illuminated by 
other than electrical means, with an initial brightness of at least 160 
microlamberts. The color may be reversed in the case of a sign that is 
self-illuminated by other than electrical means.
    (b) Floor Proximity Escape Path Marking. In lieu of the requirements 
of Sec. 25.812(e)(1), for cabin seating compartments that do not have 
the main cabin aisle entering and exiting the compartment, the following 
are applicable:
    (1) After a passenger leaves any passenger seat in the compartment, 
he/she must be able to exit the compartment to the main cabin aisle 
using only markings and visual features not more that 4 feet above the 
cabin floor, and
    (2) Proceed to the exits using the marking system necessary to 
accomplish the actions in Sec. 25.812(e)(1) and (e)(2).
    (c) Transverse Separation of the Fuselage. In the event of a 
transverse separation of the fuselage, compliance must be shown with 
Sec. 25.812(l) except as follows:
    (1) For each airplane type originally type certificated with a 
maximum passenger seating capacity of 9 or less, not more than 50 
percent of all electrically illuminated emergency lights required by 
Sec. 25.812 may be rendered inoperative in addition to the lights that 
are directly damaged by the separation.
    (2) For each airplane type originally type certificated with a 
maximum passenger seating capacity of 10 to 19, not more than 33 percent 
of all electrically illuminated emergency lights required by Sec. 
25.812 may be rendered inoperative in addition to the lights that are 
directly damaged by the separation.
    10. Interior doors. In lieu of the requirements of Sec. 25.813(e), 
interior doors may be installed between passenger seats and exits, 
provided the following requirements are met.
    (a) Each door between any passenger seat, occupiable for taxi, 
takeoff, and landing, and any emergency exit must have a means to signal 
to the flightcrew, at the flightdeck, that the door is in the open 
position for taxi, takeoff and landing.
    (b) Appropriate procedures/limitations must be established to ensure 
that any such door is in the open configuration for takeoff and landing.
    (c) Each door between any passenger seat and any exit must have dual 
means to retain it in the open position, each of which is capable of 
reacting the inertia loads specified in Sec. 25.561.
    (d) Doors installed across a longitudinal aisle must translate 
laterally to open and close, e.g., pocket doors.
    (e) Each door between any passenger seat and any exit must be 
frangible in either direction.
    (f) Each door between any passenger seat and any exit must be 
operable from either side, and if a locking mechanism is installed, it 
must be capable of being unlocked from either side without the use of 
special tools.
    11. Width of Aisle. Compliance is required with Sec. 25.815, except 
that aisle width may be reduced to 0 inches between passenger seats 
during in-flight operations only, provided that the applicant 
demonstrates that all areas of the cabin are easily accessible by a crew 
member in the event of an emergency (e.g., in-flight fire, 
decompression). Additionally, instructions must be provided at each 
passenger seat for restoring the aisle width required by Sec. 25.815. 
Procedures must be established and documented in the AFM to ensure that 
the required aisle widths are provided during taxi, takeoff, and 
landing.
    12. Materials for Compartment Interiors. Compliance is required with 
the applicable provisions of Sec. 25.853, except that compliance with 
appendix F, parts IV and V, to part 25, need not be demonstrated if it 
can be shown by test or a combination of test and analysis that the 
maximum time for evacuation of all occupants does not exceed 45 seconds 
under the conditions specified in appendix J to part 25.
    13. Fire Detection. For airplanes with a type certificated passenger 
capacity of 20 or more, there must be means that meet the requirements 
of Sec. 25.858(a) through (d) to signal the flightcrew in the event of 
a fire in any isolated room not occupiable for taxi, takeoff and 
landing, which can be closed off from the rest of the cabin by a door. 
The indication must identify the compartment where the fire is located. 
This does not apply to lavatories, which continue to be governed by 
Sec. 25.854.
    14. Cooktops. Each cooktop must be designed and installed to 
minimize any potential threat to the airplane, passengers, and crew. 
Compliance with this requirement must be found in accordance with the 
following criteria:
    (a) Means, such as conspicuous burner-on indicators, physical 
barriers, or handholds, must be installed to minimize the potential for 
inadvertent personnel contact with hot surfaces of both the cooktop and 
cookware. Conditions of turbulence must be considered.
    (b) Sufficient design means must be included to restrain cookware 
while in place on the cooktop, as well as representative contents, e.g., 
soup, sauces, etc., from the effects of flight loads and turbulence. 
Restraints must be provided to preclude hazardous movement of cookware 
and contents. These restraints must accommodate any cookware that is 
identified for use with the cooktop. Restraints must be designed to be 
easily utilized and effective in service. The cookware restraint system 
should also be designed so that it will not be easily disabled, thus 
rendering it unusable. Placarding must

[[Page 370]]

be installed which prohibits the use of cookware that cannot be 
accommodated by the restraint system.
    (c) Placarding must be installed which prohibits the use of cooktops 
(i.e., power on any burner) during taxi, takeoff, and landing.
    (d) Means must be provided to address the possibility of a fire 
occurring on or in the immediate vicinity of the cooktop. Two acceptable 
means of complying with this requirement are as follows:
    (1) Placarding must be installed that prohibits any burner from 
being powered when the cooktop is unattended. (Note: This would prohibit 
a single person from cooking on the cooktop and intermittently serving 
food to passengers while any burner is powered.) A fire detector must be 
installed in the vicinity of the cooktop which provides an audible 
warning in the passenger cabin, and a fire extinguisher of appropriate 
size and extinguishing agent must be installed in the immediate vicinity 
of the cooktop. Access to the extinguisher may not be blocked by a fire 
on or around the cooktop.
    (2) An automatic, thermally activated fire suppression system must 
be installed to extinguish a fire at the cooktop and immediately 
adjacent surfaces. The agent used in the system must be an approved 
total flooding agent suitable for use in an occupied area. The fire 
suppression system must have a manual override. The automatic activation 
of the fire suppression system must also automatically shut off power to 
the cooktop.
    (e) The surfaces of the galley surrounding the cooktop which would 
be exposed to a fire on the cooktop surface or in cookware on the 
cooktop must be constructed of materials that comply with the 
flammability requirements of part III of appendix F to part 25. This 
requirement is in addition to the flammability requirements typically 
required of the materials in these galley surfaces. During the selection 
of these materials, consideration must also be given to ensure that the 
flammability characteristics of the materials will not be adversely 
affected by the use of cleaning agents and utensils used to remove 
cooking stains.
    (f) The cooktop must be ventilated with a system independent of the 
airplane cabin and cargo ventilation system. Procedures and time 
intervals must be established to inspect and clean or replace the 
ventilation system to prevent a fire hazard from the accumulation of 
flammable oils and be included in the instructions for continued 
airworthiness. The ventilation system ducting must be protected by a 
flame arrestor. [Note: The applicant may find additional useful 
information in Society of Automotive Engineers, Aerospace Recommended 
Practice 85, Rev. E, entitled ``Air Conditioning Systems for Subsonic 
Airplanes,'' dated August 1, 1991.]
    (g) Means must be provided to contain spilled foods or fluids in a 
manner that will prevent the creation of a slipping hazard to occupants 
and will not lead to the loss of structural strength due to airplane 
corrosion.
    (h) Cooktop installations must provide adequate space for the user 
to immediately escape a hazardous cooktop condition.
    (i) A means to shut off power to the cooktop must be provided at the 
galley containing the cooktop and in the cockpit. If additional switches 
are introduced in the cockpit, revisions to smoke or fire emergency 
procedures of the AFM will be required.
    (j) If the cooktop is required to have a lid to enclose the cooktop 
there must be a means to automatically shut off power to the cooktop 
when the lid is closed.
    15. Hand-Held Fire Extinguishers.
    (a) For airplanes that were originally type certificated with more 
than 60 passengers, the number of hand-held fire extinguishers must be 
the greater of--
    (1) That provided in accordance with the requirements of Sec. 
25.851, or
    (2) A number equal to the number of originally type certificated 
exit pairs, regardless of whether the exits are deactivated for the 
proposed configuration.
    (b) Extinguishers must be evenly distributed throughout the cabin. 
These extinguishers are in addition to those required by paragraph 14 of 
this SFAR, unless it can be shown that the cooktop was installed in the 
immediate vicinity of the original exits.
    16. Security. The requirements of Sec. 25.795 are not applicable to 
airplanes approved in accordance with this SFAR.

[Doc. No. FAA-2007-28250, 74 FR 21541, May 8, 2009]



                            Subpart A_General



Sec. 25.1  Applicability.

    (a) This part prescribes airworthiness standards for the issue of 
type certificates, and changes to those certificates, for transport 
category airplanes.
    (b) Each person who applies under Part 21 for such a certificate or 
change must show compliance with the applicable requirements in this 
part.



Sec. 25.2  Special retroactive requirements.

    The following special retroactive requirements are applicable to an 
airplane for which the regulations referenced in the type certificate 
predate the sections specified below--
    (a) Irrespective of the date of application, each applicant for a 
supplemental type certificate (or an amendment to a type certificate) 
involving an increase

[[Page 371]]

in passenger seating capacity to a total greater than that for which the 
airplane has been type certificated must show that the airplane 
concerned meets the requirements of:
    (1) Sections 25.721(d), 25.783(g), 25.785(c), 25.803(c)(2) through 
(9), 25.803 (d) and (e), 25.807 (a), (c), and (d), 25.809 (f) and (h), 
25.811, 25.812, 25.813 (a), (b), and (c), 25.815, 25.817, 25.853 (a) and 
(b), 25.855(a), 25.993(f), and 25.1359(c) in effect on October 24, 1967, 
and
    (2) Sections 25.803(b) and 25.803(c)(1) in effect on April 23, 1969.
    (b) Irrespective of the date of application, each applicant for a 
supplemental type certificate (or an amendment to a type certificate) 
for an airplane manufactured after October 16, 1987, must show that the 
airplane meets the requirements of Sec. 25.807(c)(7) in effect on July 
24, 1989.
    (c) Compliance with subsequent revisions to the sections specified 
in paragraph (a) or (b) of this section may be elected or may be 
required in accordance with Sec. 21.101(a) of this chapter.

[Amdt. 25-72, 55 FR 29773, July 20, 1990, as amended by Amdt. 25-99, 65 
FR 36266, June 7, 2000]



Sec. 25.3  Special provisions for ETOPS type design approvals.

    (a) Applicability. This section applies to an applicant for ETOPS 
type design approval of an airplane:
    (1) That has an existing type certificate on February 15, 2007; or
    (2) For which an application for an original type certificate was 
submitted before February 15, 2007.
    (b) Airplanes with two engines. (1) For ETOPS type design approval 
of an airplane up to and including 180 minutes, an applicant must comply 
with Sec. 25.1535, except that it need not comply with the following 
provisions of Appendix K, K25.1.4, of this part:
    (i) K25.1.4(a), fuel system pressure and flow requirements;
    (ii) K25.1.4(a)(3), low fuel alerting; and
    (iii) K25.1.4(c), engine oil tank design.
    (2) For ETOPS type design approval of an airplane beyond 180 minutes 
an applicant must comply with Sec. 25.1535.
    (c) Airplanes with more than two engines. An applicant for ETOPS 
type design approval must comply with Sec. 25.1535 for an airplane 
manufactured on or after February 17, 2015, except that, for an airplane 
configured for a three person flight crew, the applicant need not comply 
with Appendix K, K25.1.4(a)(3), of this part, low fuel alerting.

[Doc. No. FAA-2002-6717, 72 FR 1873, Jan. 16, 2007]



Sec. 25.5  Incorporations by reference.

    (a) The materials listed in this section are incorporated by 
reference in the corresponding sections noted. These incorporations by 
reference were approved by the Director of the Federal Register in 
accordance with 5 U.S.C. 552(a) and 1 CFR part 51. These materials are 
incorporated as they exist on the date of the approval, and notice of 
any change in these materials will be published in the Federal Register. 
The materials are available for purchase at the corresponding addresses 
noted below, and all are available for inspection at the National 
Archives and Records Administration (NARA), and at FAA, Transport 
Airplane Directorate, Aircraft Certification Service, 1601 Lind Avenue, 
SW., Renton, Washington 98057-3356. For information on the availability 
of this material at NARA, call 202-741-6030, or go to: http://
www.archives.gov/federal--register/code--of--federal--regulations/ibr--
locations.html.
    (b) The following materials are available for purchase from the 
following address: The National Technical Information Services (NTIS), 
Springfield, Virginia 22166.
    (1) Fuel Tank Flammability Assessment Method User's Manual, dated 
May 2008, document number DOT/FAA/AR-05/8, IBR approved for Sec. 25.981 
and Appendix N. It can also be obtained at the following Web site: 
http://www.fire.tc.faa.gov/systems/fueltank/FTFAM.stm.
    (2) [Reserved]

[73 FR 42494, July 21, 2008]

[[Page 372]]



                            Subpart B_Flight

                                 General



Sec. 25.21  Proof of compliance.

    (a) Each requirement of this subpart must be met at each appropriate 
combination of weight and center of gravity within the range of loading 
conditions for which certification is requested. This must be shown--
    (1) By tests upon an airplane of the type for which certification is 
requested, or by calculations based on, and equal in accuracy to, the 
results of testing; and
    (2) By systematic investigation of each probable combination of 
weight and center of gravity, if compliance cannot be reasonably 
inferred from combinations investigated.
    (b) [Reserved]
    (c) The controllability, stability, trim, and stalling 
characteristics of the airplane must be shown for each altitude up to 
the maximum expected in operation.
    (d) Parameters critical for the test being conducted, such as 
weight, loading (center of gravity and inertia), airspeed, power, and 
wind, must be maintained within acceptable tolerances of the critical 
values during flight testing.
    (e) If compliance with the flight characteristics requirements is 
dependent upon a stability augmentation system or upon any other 
automatic or power-operated system, compliance must be shown with 
Sec. Sec. 25.671 and 25.672.
    (f) In meeting the requirements of Sec. Sec. 25.105(d), 25.125, 
25.233, and 25.237, the wind velocity must be measured at a height of 10 
meters above the surface, or corrected for the difference between the 
height at which the wind velocity is measured and the 10-meter height.
    (g) The requirements of this subpart associated with icing 
conditions apply only if the applicant is seeking certification for 
flight in icing conditions.
    (1) Each requirement of this subpart, except Sec. Sec. 25.121(a), 
25.123(c), 25.143(b)(1) and (2), 25.149, 25.201(c)(2), 25.239, and 
25.251(b) through (e), must be met in icing conditions. Section 
25.207(c) and (d) must be met in the landing configuration in icing 
conditions, but need not be met for other configurations. Compliance 
must be shown using the ice accretions defined in appendix C, assuming 
normal operation of the airplane and its ice protection system in 
accordance with the operating limitations and operating procedures 
established by the applicant and provided in the Airplane Flight Manual.
    (2) No changes in the load distribution limits of Sec. 25.23, the 
weight limits of Sec. 25.25 (except where limited by performance 
requirements of this subpart), and the center of gravity limits of Sec. 
25.27, from those for non-icing conditions, are allowed for flight in 
icing conditions or with ice accretion.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5671, Apr. 8, 1970; Amdt. 25-42, 43 FR 2320, Jan. 16, 1978; Amdt. 
25-72, 55 FR 29774, July 20, 1990; Amdt. 25-121, 72 FR 44665, Aug. 8, 
2007 Amdt. 25-135, 76 FR 74654, Dec. 1, 2011]

    Effective Date Note: At 79 FR 65524, Nov. 4, 2014, Sec. 25.21 was 
amended by revising paragraphs (g)(1) and (2) and adding paragraphs 
(g)(3) and (4), effective Jan. 5, 2015. For the convenience of the user, 
the added and revised text is set forth as follows:



Sec. 25.21  Proof of compliance.

                                * * * * *

    (g) * * *
    (1) Paragraphs (g)(3) and (4) of this section apply only to 
airplanes with one or both of the following attributes:
    (i) Maximum takeoff gross weight is less than 60,000 lbs; or
    (ii) The airplane is equipped with reversible flight controls.
    (2) Each requirement of this subpart, except Sec. Sec. 25.121(a), 
25.123(c), 25.143(b)(1) and (2), 25.149, 25.201(c)(2), 25.239, and 
25.251(b) through (e), must be met in the icing conditions specified in 
Appendix C of this part. Section 25.207(c) and (d) must be met in the 
landing configuration in the icing conditions specified in Appendix C, 
but need not be met for other configurations. Compliance must be shown 
using the ice accretions defined in part II of Appendix C of this part, 
assuming normal operation of the airplane and its ice protection system 
in accordance with the operating limitations and operating procedures 
established by the applicant and provided in the airplane flight manual.
    (3) If the applicant does not seek certification for flight in all 
icing conditions defined in Appendix O of this part, each requirement of 
this subpart, except Sec. Sec. 25.105, 25.107, 25.109, 25.111, 25.113, 
25.115, 25.121, 25.123, 25.143(b)(1), (b)(2), and (c)(1), 25.149, 
25.201(c)(2), 25.207(c), (d), and (e)(1), 25.239, and

[[Page 373]]

25.251(b) through (e), must be met in the Appendix O icing conditions 
for which certification is not sought in order to allow a safe exit from 
those conditions. Compliance must be shown using the ice accretions 
defined in part II, paragraphs (b) and (d) of Appendix O, assuming 
normal operation of the airplane and its ice protection system in 
accordance with the operating limitations and operating procedures 
established by the applicant and provided in the airplane flight manual.
    (4) If the applicant seeks certification for flight in any portion 
of the icing conditions of Appendix O of this part, each requirement of 
this subpart, except Sec. Sec. 25.121(a), 25.123(c), 25.143(b)(1) and 
(2), 25.149, 25.201(c)(2), 25.239, and 25.251(b) through (e), must be 
met in the Appendix O icing conditions for which certification is 
sought. Section 25.207(c) and (d) must be met in the landing 
configuration in the Appendix O icing conditions for which certification 
is sought, but need not be met for other configurations. Compliance must 
be shown using the ice accretions defined in part II, paragraphs (c) and 
(d) of Appendix O, assuming normal operation of the airplane and its ice 
protection system in accordance with the operating limitations and 
operating procedures established by the applicant and provided in the 
airplane flight manual.



Sec. 25.23  Load distribution limits.

    (a) Ranges of weights and centers of gravity within which the 
airplane may be safely operated must be established. If a weight and 
center of gravity combination is allowable only within certain load 
distribution limits (such as spanwise) that could be inadvertently 
exceeded, these limits and the corresponding weight and center of 
gravity combinations must be established.
    (b) The load distribution limits may not exceed--
    (1) The selected limits;
    (2) The limits at which the structure is proven; or
    (3) The limits at which compliance with each applicable flight 
requirement of this subpart is shown.



Sec. 25.25  Weight limits.

    (a) Maximum weights. Maximum weights corresponding to the airplane 
operating conditions (such as ramp, ground or water taxi, takeoff, en 
route, and landing), environmental conditions (such as altitude and 
temperature), and loading conditions (such as zero fuel weight, center 
of gravity position and weight distribution) must be established so that 
they are not more than--
    (1) The highest weight selected by the applicant for the particular 
conditions; or
    (2) The highest weight at which compliance with each applicable 
structural loading and flight requirement is shown, except that for 
airplanes equipped with standby power rocket engines the maximum weight 
must not be more than the highest weight established in accordance with 
appendix E of this part; or
    (3) The highest weight at which compliance is shown with the 
certification requirements of Part 36 of this chapter.
    (b) Minimum weight. The minimum weight (the lowest weight at which 
compliance with each applicable requirement of this part is shown) must 
be established so that it is not less than--
    (1) The lowest weight selected by the applicant;
    (2) The design minimum weight (the lowest weight at which compliance 
with each structural loading condition of this part is shown); or
    (3) The lowest weight at which compliance with each applicable 
flight requirement is shown.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5671, Apr. 8, 1970; Amdt. 25-63, 53 FR 16365, May 6, 1988]



Sec. 25.27  Center of gravity limits.

    The extreme forward and the extreme aft center of gravity 
limitations must be established for each practicably separable operating 
condition. No such limit may lie beyond--
    (a) The extremes selected by the applicant;
    (b) The extremes within which the structure is proven; or
    (c) The extremes within which compliance with each applicable flight 
requirement is shown.



Sec. 25.29  Empty weight and corresponding center of gravity.

    (a) The empty weight and corresponding center of gravity must be 
determined by weighing the airplane with--
    (1) Fixed ballast;
    (2) Unusable fuel determined under Sec. 25.959; and
    (3) Full operating fluids, including--
    (i) Oil;

[[Page 374]]

    (ii) Hydraulic fluid; and
    (iii) Other fluids required for normal operation of airplane 
systems, except potable water, lavatory precharge water, and fluids 
intended for injection in the engine.
    (b) The condition of the airplane at the time of determining empty 
weight must be one that is well defined and can be easily repeated.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 
43 FR 2320, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990]



Sec. 25.31  Removable ballast.

    Removable ballast may be used on showing compliance with the flight 
requirements of this subpart.



Sec. 25.33  Propeller speed and pitch limits.

    (a) The propeller speed and pitch must be limited to values that 
will ensure--
    (1) Safe operation under normal operating conditions; and
    (2) Compliance with the performance requirements of Sec. Sec. 
25.101 through 25.125.
    (b) There must be a propeller speed limiting means at the governor. 
It must limit the maximum possible governed engine speed to a value not 
exceeding the maximum allowable r.p.m.
    (c) The means used to limit the low pitch position of the propeller 
blades must be set so that the engine does not exceed 103 percent of the 
maximum allowable engine rpm or 99 percent of an approved maximum 
overspeed, whichever is greater, with--
    (1) The propeller blades at the low pitch limit and governor 
inoperative;
    (2) The airplane stationary under standard atmospheric conditions 
with no wind; and
    (3) The engines operating at the takeoff manifold pressure limit for 
reciprocating engine powered airplanes or the maximum takeoff torque 
limit for turbopropeller engine-powered airplanes.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 
49 FR 6848, Feb. 23, 1984; Amdt. 25-72, 55 FR 29774, July 20, 1990]

                               Performance



Sec. 25.101  General.

    (a) Unless otherwise prescribed, airplanes must meet the applicable 
performance requirements of this subpart for ambient atmospheric 
conditions and still air.
    (b) The performance, as affected by engine power or thrust, must be 
based on the following relative humidities;
    (1) For turbine engine powered airplanes, a relative humidity of--
    (i) 80 percent, at and below standard temperatures; and
    (ii) 34 percent, at and above standard temperatures plus 50 [deg]F.

Between these two temperatures, the relative humidity must vary 
linearly.
    (2) For reciprocating engine powered airplanes, a relative humidity 
of 80 percent in a standard atmosphere. Engine power corrections for 
vapor pressure must be made in accordance with the following table:

------------------------------------------------------------------------
                               Specific humidity w
 Altitude H    Vapor pressure   (Lb. moisture per   Density ratio [rho]
    (ft.)       e (In. Hg.)        lb. dry air)       [sigma]=0.0023769
------------------------------------------------------------------------
         0            0.403              0.00849               0.99508
     1,000             .354               .00773                .96672
     2,000             .311               .00703                .93895
     3,000             .272               .00638                .91178
     4,000             .238               .00578                .88514
     5,000             .207               .00523                .85910
     6,000            .1805               .00472                .83361
     7,000            .1566               .00425                .80870
     8,000            .1356               .00382                .78434
     9,000            .1172               .00343                .76053
    10,000            .1010               .00307                .73722
    15,000            .0463              .001710                .62868
    20,000           .01978              .000896                .53263
    25,000           .00778              .000436                .44806
------------------------------------------------------------------------

    (c) The performance must correspond to the propulsive thrust 
available under the particular ambient atmospheric conditions, the 
particular flight condition, and the relative humidity specified in 
paragraph (b) of this section. The available propulsive thrust must 
correspond to engine power or thrust, not exceeding the approved power 
or thrust less--
    (1) Installation losses; and
    (2) The power or equivalent thrust absorbed by the accessories and 
services appropriate to the particular ambient atmospheric conditions 
and the particular flight condition.
    (d) Unless otherwise prescribed, the applicant must select the 
takeoff, en route, approach, and landing configurations for the 
airplane.
    (e) The airplane configurations may vary with weight, altitude, and 
temperature, to the extent they are compatible with the operating 
procedures

[[Page 375]]

required by paragraph (f) of this section.
    (f) Unless otherwise prescribed, in determining the accelerate-stop 
distances, takeoff flight paths, takeoff distances, and landing 
distances, changes in the airplane's configuration, speed, power, and 
thrust, must be made in accordance with procedures established by the 
applicant for operation in service.
    (g) Procedures for the execution of balked landings and missed 
approaches associated with the conditions prescribed in Sec. Sec. 
25.119 and 25.121(d) must be established.
    (h) The procedures established under paragraphs (f) and (g) of this 
section must--
    (1) Be able to be consistently executed in service by crews of 
average skill;
    (2) Use methods or devices that are safe and reliable; and
    (3) Include allowance for any time delays, in the execution of the 
procedures, that may reasonably be expected in service.
    (i) The accelerate-stop and landing distances prescribed in 
Sec. Sec. 25.109 and 25.125, respectively, must be determined with all 
the airplane wheel brake assemblies at the fully worn limit of their 
allowable wear range.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55466, Dec. 20, 1976; Amdt. 25-92, 63 FR 8318, Feb. 18, 1998]



Sec. 25.103  Stall speed.

    (a) The reference stall speed, VSR, is a calibrated 
airspeed defined by the applicant. VSR may not be less than a 
1-g stall speed. VSR is expressed as:
[GRAPHIC] [TIFF OMITTED] TR26NO02.001

where:
VCLMAX = Calibrated airspeed obtained when the load factor-
          corrected lift coefficient
          [GRAPHIC] [TIFF OMITTED] TR26NO02.002
          
    is first a maximum during the maneuver prescribed in paragraph (c) 
of this section. In addition, when the maneuver is limited by a device 
that abruptly pushes the nose down at a selected angle of attack (e.g., 
a stick pusher), VCLMAX may not be less than the speed 
existing at the instant the device operates;
nZW = Load factor normal to the flight path at 
          VCLMAX
W = Airplane gross weight;
S = Aerodynamic reference wing area; and
q = Dynamic pressure.

    (b) VCLMAX is determined with:
    (1) Engines idling, or, if that resultant thrust causes an 
appreciable decrease in stall speed, not more than zero thrust at the 
stall speed;
    (2) Propeller pitch controls (if applicable) in the takeoff 
position;
    (3) The airplane in other respects (such as flaps, landing gear, and 
ice accretions) in the condition existing in the test or performance 
standard in which VSR is being used;
    (4) The weight used when VSR is being used as a factor to 
determine compliance with a required performance standard;
    (5) The center of gravity position that results in the highest value 
of reference stall speed; and
    (6) The airplane trimmed for straight flight at a speed selected by 
the applicant, but not less than 1.13VSR and not greater than 
1.3VSR.
    (c) Starting from the stabilized trim condition, apply the 
longitudinal control to decelerate the airplane so that the speed 
reduction does not exceed one knot per second.
    (d) In addition to the requirements of paragraph (a) of this 
section, when a device that abruptly pushes the nose down at a selected 
angle of attack (e.g., a stick pusher) is installed, the reference stall 
speed, VSR, may not be less than 2 knots or 2 percent, 
whichever is greater, above the speed at which the device operates.

[Doc. No. 28404, 67 FR 70825, Nov. 26, 2002, as amended by Amdt. 25-121, 
72 FR 44665, Aug. 8, 2007]



Sec. 25.105  Takeoff.

    (a) The takeoff speeds prescribed by Sec. 25.107, the accelerate-
stop distance prescribed by Sec. 25.109, the takeoff path prescribed by 
Sec. 25.111, the takeoff distance and takeoff run prescribed by Sec. 
25.113, and the net takeoff flight path prescribed by Sec. 25.115, must 
be determined in the selected configuration for takeoff at each weight, 
altitude, and

[[Page 376]]

ambient temperature within the operational limits selected by the 
applicant--
    (1) In non-icing conditions; and
    (2) In icing conditions, if in the configuration of Sec. 25.121(b) 
with the takeoff ice accretion defined in appendix C:
    (i) The stall speed at maximum takeoff weight exceeds that in non-
icing conditions by more than the greater of 3 knots CAS or 3 percent of 
VSR; or
    (ii) The degradation of the gradient of climb determined in 
accordance with Sec. 25.121(b) is greater than one-half of the 
applicable actual-to-net takeoff flight path gradient reduction defined 
in Sec. 25.115(b).
    (b) No takeoff made to determine the data required by this section 
may require exceptional piloting skill or alertness.
    (c) The takeoff data must be based on--
    (1) In the case of land planes and amphibians:
    (i) Smooth, dry and wet, hard-surfaced runways; and
    (ii) At the option of the applicant, grooved or porous friction 
course wet, hard-surfaced runways.
    (2) Smooth water, in the case of seaplanes and amphibians; and
    (3) Smooth, dry snow, in the case of skiplanes.
    (d) The takeoff data must include, within the established 
operational limits of the airplane, the following operational correction 
factors:
    (1) Not more than 50 percent of nominal wind components along the 
takeoff path opposite to the direction of takeoff, and not less than 150 
percent of nominal wind components along the takeoff path in the 
direction of takeoff.
    (2) Effective runway gradients.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-92, 
63 FR 8318, Feb. 18, 1998; Amdt. 25-121, 72 FR 44665, Aug. 8, 2007]

    Effective Date Note: At 79 FR 65525, Nov. 4, 2014, Sec. 25.105 was 
amended by revising paragraph (a)(2) introductory text, effective Jan. 
5, 2015. For the convenience of the user, the revised text is set forth 
as follows:



Sec. 25.105  Takeoff.

    (a) * * *
    (2) In icing conditions, if in the configuration used to show 
compliance with Sec. 25.121(b), and with the most critical of the 
takeoff ice accretion(s) defined in Appendices C and O of this part, as 
applicable, in accordance with Sec. 25.21(g):

                                * * * * *



Sec. 25.107  Takeoff speeds.

    (a) V1 must be established in relation to VEF as follows:
    (1) VEF is the calibrated airspeed at which the critical engine is 
assumed to fail. VEF must be selected by the applicant, but may not be 
less than VMCG determined under Sec. 25.149(e).
    (2) V1, in terms of calibrated airspeed, is selected by 
the applicant; however, V1 may not be less than 
VEF plus the speed gained with critical engine inoperative 
during the time interval between the instant at which the critical 
engine is failed, and the instant at which the pilot recognizes and 
reacts to the engine failure, as indicated by the pilot's initiation of 
the first action (e.g., applying brakes, reducing thrust, deploying 
speed brakes) to stop the airplane during accelerate-stop tests.
    (b) V2MIN, in terms of calibrated airspeed, may not be less than--
    (1) 1.13 VSR for--
    (i) Two-engine and three-engine turbopropeller and reciprocating 
engine powered airplanes; and
    (ii) Turbojet powered airplanes without provisions for obtaining a 
significant reduction in the one-engine-inoperative power-on stall 
speed;
    (2) 1.08 VSR for--
    (i) Turbopropeller and reciprocating engine powered airplanes with 
more than three engines; and
    (ii) Turbojet powered airplanes with provisions for obtaining a 
significant reduction in the one-engine-inoperative power-on stall 
speed; and
    (3) 1.10 times VMC established under Sec. 25.149.
    (c) V2, in terms of calibrated airspeed, must be selected by the 
applicant to provide at least the gradient of climb required by Sec. 
25.121(b) but may not be less than--
    (1) V2MIN;
    (2) VR plus the speed increment attained (in accordance with Sec. 
25.111(c)(2)) before reaching a height of 35 feet above the takeoff 
surface; and
    (3) A speed that provides the maneuvering capability specified in 
Sec. 25.143(h).

[[Page 377]]

    (d) VMU is the calibrated airspeed at and above which the airplane 
can safely lift off the ground, and con- tinue the takeoff. VMU speeds 
must be selected by the applicant throughout the range of thrust-to-
weight ratios to be certificated. These speeds may be established from 
free air data if these data are verified by ground takeoff tests.
    (e) VR, in terms of calibrated airspeed, must be selected in 
accordance with the conditions of paragraphs (e)(1) through (4) of this 
section:
    (1) VR may not be less than--
    (i) V1;
    (ii) 105 percent of VMC;
    (iii) The speed (determined in accordance with Sec. 25.111(c)(2)) 
that allows reaching V2 before reaching a height of 35 feet 
above the takeoff surface; or
    (iv) A speed that, if the airplane is rotated at its maximum 
practicable rate, will result in a VLOF of not less than --
    (A) 110 percent of VMU in the all-engines-operating 
condition, and 105 percent of VMU determined at the thrust-
to-weight ratio corresponding to the one-engine-inoperative condition; 
or
    (B) If the VMU attitude is limited by the geometry of the 
airplane (i.e., tail contact with the runway), 108 percent of 
VMU in the all-engines-operating condition, and 104 percent 
of VMU determined at the thrust-to-weight ratio corresponding 
to the one-engine-inoperative condition.
    (2) For any given set of conditions (such as weight, configuration, 
and temperature), a single value of VR, obtained in accordance with this 
paragraph, must be used to show compliance with both the one-engine-
inoperative and the all-engines-operating takeoff provisions.
    (3) It must be shown that the one-engine-inoperative takeoff 
distance, using a rotation speed of 5 knots less than VR established in 
accordance with paragraphs (e)(1) and (2) of this section, does not 
exceed the corresponding one-engine-inoperative takeoff distance using 
the established VR. The takeoff distances must be determined in 
accordance with Sec. 25.113(a)(1).
    (4) Reasonably expected variations in service from the established 
takeoff procedures for the operation of the airplane (such as over-
rotation of the airplane and out-of-trim conditions) may not result in 
unsafe flight characteristics or in marked increases in the scheduled 
takeoff distances established in accordance with Sec. 25.113(a).
    (f) VLOF is the calibrated airspeed at which the airplane first 
becomes airborne.
    (g) VFTO, in terms of calibrated airspeed, must be 
selected by the applicant to provide at least the gradient of climb 
required by Sec. 25.121(c), but may not be less than--
    (1) 1.18 VSR; and
    (2) A speed that provides the maneuvering capability specified in 
Sec. 25.143(h).
    (h) In determining the takeoff speeds V1, VR, 
and V2 for flight in icing conditions, the values of 
VMCG, VMC, and VMU determined for non-
icing conditions may be used.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55466, Dec. 20, 1976; Amdt. 25-42, 43 FR 2320, Jan. 16, 1978; 
Amdt. 25-92, 63 FR 8318, Feb. 18, 1998; Amdt. 25-94, 63 FR 8848, Feb. 
23, 1998; Amdt. 25-108, 67 FR 70826, Nov. 26, 2002; Amdt. 25-121, 72 FR 
44665, Aug. 8, 2007; Amdt. 25-135, 76 FR 74654, Dec. 1, 2011]



Sec. 25.109  Accelerate-stop distance.

    (a) The accelerate-stop distance on a dry runway is the greater of 
the following distances:
    (1) The sum of the distances necessary to--
    (i) Accelerate the airplane from a standing start with all engines 
operating to VEF for takeoff from a dry runway;
    (ii) Allow the airplane to accelerate from VEF to the 
highest speed reached during the rejected takeoff, assuming the critical 
engine fails at VEF and the pilot takes the first action to 
reject the takeoff at the V1 for takeoff from a dry runway; 
and
    (iii) Come to a full stop on a dry runway from the speed reached as 
prescribed in paragraph (a)(1)(ii) of this section; plus
    (iv) A distance equivalent to 2 seconds at the V1 for 
takeoff from a dry runway.
    (2) The sum of the distances necessary to--

[[Page 378]]

    (i) Accelerate the airplane from a standing start with all engines 
operating to the highest speed reached during the rejected takeoff, 
assuming the pilot takes the first action to reject the takeoff at the 
V1 for takeoff from a dry runway; and
    (ii) With all engines still operating, come to a full stop on dry 
runway from the speed reached as prescribed in paragraph (a)(2)(i) of 
this section; plus
    (iii) A distance equivalent to 2 seconds at the V1 for 
takeoff from a dry runway.
    (b) The accelerate-stop distance on a wet runway is the greater of 
the following distances:
    (1) The accelerate-stop distance on a dry runway determined in 
accordance with paragraph (a) of this section; or
    (2) The accelerate-stop distance determined in accordance with 
paragraph (a) of this section, except that the runway is wet and the 
corresponding wet runway values of VEF and V1 are 
used. In determining the wet runway accelerate-stop distance, the 
stopping force from the wheel brakes may never exceed:
    (i) The wheel brakes stopping force determined in meeting the 
requirements of Sec. 25.101(i) and paragraph (a) of this section; and
    (ii) The force resulting from the wet runway braking coefficient of 
friction determined in accordance with paragraphs (c) or (d) of this 
section, as applicable, taking into account the distribution of the 
normal load between braked and unbraked wheels at the most adverse 
center-of-gravity position approved for takeoff.
    (c) The wet runway braking coefficient of friction for a smooth wet 
runway is defined as a curve of friction coefficient versus ground speed 
and must be computed as follows:
    (1) The maximum tire-to-ground wet runway braking coefficient of 
friction is defined as:
[GRAPHIC] [TIFF OMITTED] TR18FE98.004

Where--

Tire Pressure=maximum airplane operating tire pressure (psi);
[mu]t/gMAX=maximum tire-to-ground braking coefficient;
V=airplane true ground speed (knots); and
Linear interpolation may be used for tire pressures other than those 
          listed.

    (2) The maximum tire-to-ground wet runway braking coefficient of 
friction must be adjusted to take into account the efficiency of the 
anti-skid system on a wet runway. Anti-skid system operation must be 
demonstrated by flight testing on a smooth wet runway, and its 
efficiency must be determined. Unless a specific anti-skid system 
efficiency is determined from a quantitative analysis of the flight 
testing on a smooth wet runway, the maximum tire-to-ground wet runway 
braking coefficient of friction determined in paragraph (c)(1) of this 
section must be multiplied by the efficiency value associated with the 
type of anti-skid system installed on the airplane:

[[Page 379]]



------------------------------------------------------------------------
                                                              Efficiency
                  Type of anti-skid system                       value
------------------------------------------------------------------------
On-Off......................................................       0.30
Quasi-Modulating............................................       0.50
Fully Modulating............................................       0.80
------------------------------------------------------------------------

    (d) At the option of the applicant, a higher wet runway braking 
coefficient of friction may be used for runway surfaces that have been 
grooved or treated with a porous friction course material. For grooved 
and porous friction course runways, the wet runway braking coefficent of 
friction is defined as either:
    (1) 70 percent of the dry runway braking coefficient of friction 
used to determine the dry runway accelerate-stop distance; or
    (2) The wet runway braking coefficient defined in paragraph (c) of 
this section, except that a specific anti-skid system efficiency, if 
determined, is appropriate for a grooved or porous friction course wet 
runway, and the maximum tire-to-ground wet runway braking coefficient of 
friction is defined as:
[GRAPHIC] [TIFF OMITTED] TR18FE98.005

Where--

Tire Pressure=maximum airplane operating tire pressure (psi);
[mu]t/gMAX=maximum tire-to-ground braking coefficient;
V=airplane true ground speed (knots); and
Linear interpolation may be used for tire pressures other than those 
          listed.

    (e) Except as provided in paragraph (f)(1) of this section, means 
other than wheel brakes may be used to determine the accelerate-stop 
distance if that means--
    (1) Is safe and reliable;
    (2) Is used so that consistent results can be expected under normal 
operating conditions; and
    (3) Is such that exceptional skill is not required to control the 
airplane.
    (f) The effects of available reverse thrust--
    (1) Shall not be included as an additional means of deceleration 
when determining the accelerate-stop distance on a dry runway; and
    (2) May be included as an additional means of deceleration using 
recommended reverse thrust procedures when determining the accelerate-
stop distance on a wet runway, provided the requirements of paragraph 
(e) of this section are met.
    (g) The landing gear must remain extended throughout the accelerate-
stop distance.
    (h) If the accelerate-stop distance includes a stopway with surface 
characteristics substantially different from those of the runway, the 
takeoff data must include operational correction factors for the 
accelerate-stop distance. The correction factors must account for the 
particular surface characteristics of the stopway and the variations in 
these characteristics with seasonal weather conditions (such as 
temperature, rain, snow, and ice) within the established operational 
limits.
    (i) A flight test demonstration of the maximum brake kinetic energy 
accelerate-stop distance must be conducted with not more than 10 percent 
of the

[[Page 380]]

allowable brake wear range remaining on each of the airplane wheel 
brakes.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 
43 FR 2321, Jan. 16, 1978; Amdt. 25-92, 63 FR 8318, Feb. 18, 1998]



Sec. 25.111  Takeoff path.

    (a) The takeoff path extends from a standing start to a point in the 
takeoff at which the airplane is 1,500 feet above the takeoff surface, 
or at which the transition from the takeoff to the en route 
configuration is completed and VFTO is reached, whichever 
point is higher. In addition--
    (1) The takeoff path must be based on the procedures prescribed in 
Sec. 25.101(f);
    (2) The airplane must be accelerated on the ground to VEF, at which 
point the critical engine must be made inoperative and remain 
inoperative for the rest of the takeoff; and
    (3) After reaching VEF, the airplane must be accelerated to 
V2.
    (b) During the acceleration to speed V2, the nose gear 
may be raised off the ground at a speed not less than VR. However, 
landing gear retraction may not be begun until the airplane is airborne.
    (c) During the takeoff path determination in accordance with 
paragraphs (a) and (b) of this section--
    (1) The slope of the airborne part of the takeoff path must be 
positive at each point;
    (2) The airplane must reach V2 before it is 35 feet above the 
takeoff surface and must continue at a speed as close as practical to, 
but not less than V2, until it is 400 feet above the takeoff surface;
    (3) At each point along the takeoff path, starting at the point at 
which the airplane reaches 400 feet above the takeoff surface, the 
available gradient of climb may not be less than--
    (i) 1.2 percent for two-engine airplanes;
    (ii) 1.5 percent for three-engine airplanes; and
    (iii) 1.7 percent for four-engine airplanes.
    (4) The airplane configuration may not be changed, except for gear 
retraction and automatic propeller feathering, and no change in power or 
thrust that requires action by the pilot may be made until the airplane 
is 400 feet above the takeoff surface; and
    (5) If Sec. 25.105(a)(2) requires the takeoff path to be determined 
for flight in icing conditions, the airborne part of the takeoff must be 
based on the airplane drag:
    (i) With the takeoff ice accretion defined in appendix C, from a 
height of 35 feet above the takeoff surface up to the point where the 
airplane is 400 feet above the takeoff surface; and
    (ii) With the final takeoff ice accretion defined in appendix C, 
from the point where the airplane is 400 feet above the takeoff surface 
to the end of the takeoff path.
    (d) The takeoff path must be determined by a continuous demonstrated 
takeoff or by synthesis from segments. If the takeoff path is determined 
by the segmental method--
    (1) The segments must be clearly defined and must be related to the 
distinct changes in the configuration, power or thrust, and speed;
    (2) The weight of the airplane, the configuration, and the power or 
thrust must be constant throughout each segment and must correspond to 
the most critical condition prevailing in the segment;
    (3) The flight path must be based on the airplane's performance 
without ground effect; and
    (4) The takeoff path data must be checked by continuous demonstrated 
takeoffs up to the point at which the airplane is out of ground effect 
and its speed is stabilized, to ensure that the path is conservative 
relative to the continous path.

The airplane is considered to be out of the ground effect when it 
reaches a height equal to its wing span.
    (e) For airplanes equipped with standby power rocket engines, the 
takeoff path may be determined in accordance with section II of appendix 
E.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-6, 30 
FR 8468, July 2, 1965; Amdt. 25-42, 43 FR 2321, Jan. 16, 1978; Amdt. 25-
54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-72, 55 FR 29774, July 20, 
1990; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998; Amdt. 25-108, 67 FR 70826, 
Nov. 26, 2002; Amdt. 25-115, 69 FR 40527, July 2, 2004; Amdt. 25-121, 72 
FR 44666; Aug. 8, 2007]

[[Page 381]]


    Effective Date Note: At 79 FR 65525, Nov. 4, 2014, Sec. 25.111 was 
amended by revising paragraphs (c)(5)(i) and (ii), effective Jan. 5, 
2015. For the convenience of the user, the revised text is set forth as 
follows:



Sec. 25.111  Takeoff path.

                                * * * * *

    (c) * * *
    (5) * * *
    (i) With the most critical of the takeoff ice accretion(s) defined 
in Appendices C and O of this part, as applicable, in accordance with 
Sec. 25.21(g), from a height of 35 feet above the takeoff surface up to 
the point where the airplane is 400 feet above the takeoff surface; and
    (ii) With the most critical of the final takeoff ice accretion(s) 
defined in Appendices C and O of this part, as applicable, in accordance 
with Sec. 25.21(g), from the point where the airplane is 400 feet above 
the takeoff surface to the end of the takeoff path.

                                * * * * *



Sec. 25.113  Takeoff distance and takeoff run.

    (a) Takeoff distance on a dry runway is the greater of--
    (1) The horizontal distance along the takeoff path from the start of 
the takeoff to the point at which the airplane is 35 feet above the 
takeoff surface, determined under Sec. 25.111 for a dry runway; or
    (2) 115 percent of the horizontal distance along the takeoff path, 
with all engines operating, from the start of the takeoff to the point 
at which the airplane is 35 feet above the takeoff surface, as 
determined by a procedure consistent with Sec. 25.111.
    (b) Takeoff distance on a wet runway is the greater of--
    (1) The takeoff distance on a dry runway determined in accordance 
with paragraph (a) of this section; or
    (2) The horizontal distance along the takeoff path from the start of 
the takeoff to the point at which the airplane is 15 feet above the 
takeoff surface, achieved in a manner consistent with the achievement of 
V2 before reaching 35 feet above the takeoff surface, 
determined under Sec. 25.111 for a wet runway.
    (c) If the takeoff distance does not include a clearway, the takeoff 
run is equal to the takeoff distance. If the takeoff distance includes a 
clearway--
    (1) The takeoff run on a dry runway is the greater of--
    (i) The horizontal distance along the takeoff path from the start of 
the takeoff to a point equidistant between the point at which 
VLOF is reached and the point at which the airplane is 35 
feet above the takeoff surface, as determined under Sec. 25.111 for a 
dry runway; or
    (ii) 115 percent of the horizontal distance along the takeoff path, 
with all engines operating, from the start of the takeoff to a point 
equidistant between the point at which VLOF is reached and 
the point at which the airplane is 35 feet above the takeoff surface, 
determined by a procedure consistent with Sec. 25.111.
    (2) The takeoff run on a wet runway is the greater of--
    (i) The horizontal distance along the takeoff path from the start of 
the takeoff to the point at which the airplane is 15 feet above the 
takeoff surface, achieved in a manner consistent with the achievement of 
V2 before reaching 35 feet above the takeoff surface, as 
determined under Sec. 25.111 for a wet runway; or
    (ii) 115 percent of the horizontal distance along the takeoff path, 
with all engines operating, from the start of the takeoff to a point 
equidistant between the point at which VLOF is reached and 
the point at which the airplane is 35 feet above the takeoff surface, 
determined by a procedure consistent with Sec. 25.111.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5671, Apr. 8, 1970; Amdt. 25-92, 63 FR 8320, Feb. 18, 1998]



Sec. 25.115  Takeoff flight path.

    (a) The takeoff flight path shall be considered to begin 35 feet 
above the takeoff surface at the end of the takeoff distance determined 
in accordance with Sec. 25.113(a) or (b), as appropriate for the runway 
surface condition.
    (b) The net takeoff flight path data must be determined so that they 
represent the actual takeoff flight paths (determined in accordance with 
Sec. 25.111 and with paragraph (a) of this section) reduced at each 
point by a gradient of climb equal to--
    (1) 0.8 percent for two-engine airplanes;

[[Page 382]]

    (2) 0.9 percent for three-engine airplanes; and
    (3) 1.0 percent for four-engine airplanes.
    (c) The prescribed reduction in climb gradient may be applied as an 
equivalent reduction in acceleration along that part of the takeoff 
flight path at which the airplane is accelerated in level flight.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-92, 
63 FR 8320, Feb. 18, 1998]



Sec. 25.117  Climb: general.

    Compliance with the requirements of Sec. Sec. 25.119 and 25.121 
must be shown at each weight, altitude, and ambient temperature within 
the operational limits established for the airplane and with the most 
unfavorable center of gravity for each configuration.



Sec. 25.119  Landing climb: All-engines-operating.

    In the landing configuration, the steady gradient of climb may not 
be less than 3.2 percent, with the engines at the power or thrust that 
is available 8 seconds after initiation of movement of the power or 
thrust controls from the minimum flight idle to the go-around power or 
thrust setting--
    (a) In non-icing conditions, with a climb speed of VREF 
determined in accordance with Sec. 25.125(b)(2)(i); and
    (b) In icing conditions with the landing ice accretion defined in 
appendix C, and with a climb speed of VREF determined in 
accordance with Sec. 25.125(b)(2)(ii).

[Amdt. 25-121, 72 FR 44666; Aug. 8, 2007]

    Effective Date Note: At 79 FR 65525, Nov. 4, 2014, Sec. 25.119 was 
amended by revising paragraph (b), effective Jan. 5, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.119  Landing climb: All-engines-operating.

                                * * * * *

    (b) In icing conditions with the most critical of the landing ice 
accretion(s) defined in Appendices C and O of this part, as applicable, 
in accordance with Sec. 25.21(g), and with a climb speed of 
VREF determined in accordance with Sec. 25.125(b)(2)(ii).



Sec. 25.121  Climb: One-engine-inoperative.

    (a) Takeoff; landing gear extended. In the critical takeoff 
configuration existing along the flight path (between the points at 
which the airplane reaches VLOF and at which the landing gear is fully 
retracted) and in the configuration used in Sec. 25.111 but without 
ground effect, the steady gradient of climb must be positive for two-
engine airplanes, and not less than 0.3 percent for three-engine 
airplanes or 0.5 percent for four-engine airplanes, at VLOF and with--
    (1) The critical engine inoperative and the remaining engines at the 
power or thrust available when retraction of the landing gear is begun 
in accordance with Sec. 25.111 unless there is a more critical power 
operating condition existing later along the flight path but before the 
point at which the landing gear is fully retracted; and
    (2) The weight equal to the weight existing when retraction of the 
landing gear is begun, determined under Sec. 25.111.
    (b) Takeoff; landing gear retracted. In the takeoff configuration 
existing at the point of the flight path at which the landing gear is 
fully retracted, and in the configuration used in Sec. 25.111 but 
without ground effect:
    (1) The steady gradient of climb may not be less than 2.4 percent 
for two-engine airplanes, 2.7 percent for three-engine airplanes, and 
3.0 percent for four-engine airplanes, at V2 with:
    (i) The critical engine inoperative, the remaining engines at the 
takeoff power or thrust available at the time the landing gear is fully 
retracted, determined under Sec. 25.111, unless there is a more 
critical power operating condition existing later along the flight path 
but before the point where the airplane reaches a height of 400 feet 
above the takeoff surface; and
    (ii) The weight equal to the weight existing when the airplane's 
landing gear is fully retracted, determined under Sec. 25.111.
    (2) The requirements of paragraph (b)(1) of this section must be 
met:
    (i) In non-icing conditions; and
    (ii) In icing conditions with the takeoff ice accretion defined in 
appendix C,

[[Page 383]]

if in the configuration of Sec. 25.121(b) with the takeoff ice 
accretion:
    (A) The stall speed at maximum takeoff weight exceeds that in non-
icing conditions by more than the greater of 3 knots CAS or 3 percent of 
VSR; or
    (B) The degradation of the gradient of climb determined in 
accordance with Sec. 25.121(b) is greater than one-half of the 
applicable actual-to-net takeoff flight path gradient reduction defined 
in Sec. 25.115(b).
    (c) Final takeoff. In the en route configuration at the end of the 
takeoff path determined in accordance with Sec. 25.111:
    (1) The steady gradient of climb may not be less than 1.2 percent 
for two-engine airplanes, 1.5 percent for three-engine airplanes, and 
1.7 percent for four-engine airplanes, at VFTO with--
    (i) The critical engine inoperative and the remaining engines at the 
available maximum continuous power or thrust; and
    (ii) The weight equal to the weight existing at the end of the 
takeoff path, determined under Sec. 25.111.
    (2) The requirements of paragraph (c)(1) of this section must be 
met:
    (i) In non-icing conditions; and
    (ii) In icing conditions with the final takeoff ice accretion 
defined in appendix C, if in the configuration of Sec. 25.121(b) with 
the takeoff ice accretion:
    (A) The stall speed at maximum takeoff weight exceeds that in non-
icing conditions by more than the greater of 3 knots CAS or 3 percent of 
VSR; or
    (B) The degradation of the gradient of climb determined in 
accordance with Sec. 25.121(b) is greater than one-half of the 
applicable actual-to-net takeoff flight path gradient reduction defined 
in Sec. 25.115(b).
    (d) Approach. In a configuration corresponding to the normal all-
engines-operating procedure in which VSR for this 
configuration does not exceed 110 percent of the VSR for the 
related all-engines-operating landing configuration:
    (1) The steady gradient of climb may not be less than 2.1 percent 
for two-engine airplanes, 2.4 percent for three-engine airplanes, and 
2.7 percent for four-engine airplanes, with--
    (i) The critical engine inoperative, the remaining engines at the 
go-around power or thrust setting;
    (ii) The maximum landing weight;
    (iii) A climb speed established in connection with normal landing 
procedures, but not exceeding 1.4 VSR; and
    (iv) Landing gear retracted.
    (2) The requirements of paragraph (d)(1) of this section must be 
met:
    (i) In non-icing conditions; and
    (ii) In icing conditions with the approach ice accretion defined in 
appendix C. The climb speed selected for non-icing conditions may be 
used if the climb speed for icing conditions, computed in accordance 
with paragraph (d)(1)(iii) of this section, does not exceed that for 
non-icing conditions by more than the greater of 3 knots CAS or 3 
percent.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-84, 
60 FR 30749, June 9, 1995; Amdt. 25-108, 67 FR 70826, Nov. 26, 2002; 
Amdt. 25-121, 72 FR 44666; Aug. 8, 2007]

    Effective Date Note: At 79 FR 65525, Nov. 4, 2014, Sec. 25.121 was 
amended by revising paragraphs (b)(2)(ii) introductory text, (c)(2)(ii) 
introductory text, and (d)(2)(ii), effective Jan. 5, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.121  Climb: One-engine-inoperative.

                                * * * * *

    (b) * * *
    (2) * * *
    (ii) In icing conditions with the most critical of the takeoff ice 
accretion(s) defined in Appendices C and O of this part, as applicable, 
in accordance with Sec. 25.21(g), if in the configuration used to show 
compliance with Sec. 25.121(b) with this takeoff ice accretion:

                                * * * * *

    (c) * * *
    (2) * * *
    (ii) In icing conditions with the most critical of the final takeoff 
ice accretion(s) defined in Appendices C and O of this part, as 
applicable, in accordance with Sec. 25.21(g), if in the configuration 
used to show compliance with Sec. 25.121(b) with the takeoff ice 
accretion used to show compliance with Sec. 25.111(c)(5)(i):

                                * * * * *

    (d) * * *
    (2) * * *

[[Page 384]]

    (ii) In icing conditions with the most critical of the approach ice 
accretion(s) defined in Appendices C and O of this part, as applicable, 
in accordance with Sec. 25.21(g). The climb speed selected for non-
icing conditions may be used if the climb speed for icing conditions, 
computed in accordance with paragraph (d)(1)(iii) of this section, does 
not exceed that for non-icing conditions by more than the greater of 3 
knots CAS or 3 percent.



Sec. 25.123  En route flight paths.

    (a) For the en route configuration, the flight paths prescribed in 
paragraph (b) and (c) of this section must be determined at each weight, 
altitude, and ambient temperature, within the operating limits 
established for the airplane. The variation of weight along the flight 
path, accounting for the progressive consumption of fuel and oil by the 
operating engines, may be included in the computation. The flight paths 
must be determined at a speed not less than VFTO, with--
    (1) The most unfavorable center of gravity;
    (2) The critical engines inoperative;
    (3) The remaining engines at the available maximum continuous power 
or thrust; and
    (4) The means for controlling the engine-cooling air supply in the 
position that provides adequate cooling in the hot-day condition.
    (b) The one-engine-inoperative net flight path data must represent 
the actual climb performance diminished by a gradient of climb of 1.1 
percent for two-engine airplanes, 1.4 percent for three-engine 
airplanes, and 1.6 percent for four-engine airplanes--
    (1) In non-icing conditions; and
    (2) In icing conditions with the en route ice accretion defined in 
appendix C, if:
    (i) A speed of 1.18 ``VSR0 with the en route ice 
accretion exceeds the en route speed selected for non-icing conditions 
by more than the greater of 3 knots CAS or 3 percent of VSR; 
or
    (ii) The degradation of the gradient of climb is greater than one-
half of the applicable actual-to-net flight path reduction defined in 
paragraph (b) of this section.
    (c) For three- or four-engine airplanes, the two-engine-inoperative 
net flight path data must represent the actual climb performance 
diminished by a gradient of climb of 0.3 percent for three-engine 
airplanes and 0.5 percent for four-engine airplanes.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-121, 
72 FR 44666; Aug. 8, 2007]

    Effective Date Note: At 79 FR 65525, Nov. 4, 2014, Sec. 25.123 was 
amended by revising paragraph (b)(2) introductory text, effective Jan. 
5, 2015. For the convenience of the user, the revised text is set forth 
as follows:



Sec. 25.123  En route flight paths.

                                * * * * *

    (b) * * *
    (2) In icing conditions with the most critical of the en route ice 
accretion(s) defined in Appendices C and O of this part, as applicable, 
in accordance with Sec. 25.21(g), if:

                                * * * * *



Sec. 25.125  Landing.

    (a) The horizontal distance necessary to land and to come to a 
complete stop (or to a speed of approximately 3 knots for water 
landings) from a point 50 feet above the landing surface must be 
determined (for standard temperatures, at each weight, altitude, and 
wind within the operational limits established by the applicant for the 
airplane):
    (1) In non-icing conditions; and
    (2) In icing conditions with the landing ice accretion defined in 
appendix C if VREF for icing conditions exceeds 
VREF for non-icing conditions by more than 5 knots CAS at the 
maximum landing weight.
    (b) In determining the distance in paragraph (a) of this section:
    (1) The airplane must be in the landing configuration.
    (2) A stabilized approach, with a calibrated airspeed of not less 
than VREF, must be maintained down to the 50-foot height.
    (i) In non-icing conditions, VREF may not be less than:
    (A) 1.23 VSR0;
    (B) VMCL established under Sec. 25.149(f); and
    (C) A speed that provides the maneuvering capability specified in 
Sec. 25.143(h).
    (ii) In icing conditions, VREF may not be less than:

[[Page 385]]

    (A) The speed determined in paragraph (b)(2)(i) of this section;
    (B) 1.23 VSR0 with the landing ice accretion defined in 
appendix C if that speed exceeds VREF for non-icing 
conditions by more than 5 knots CAS; and
    (C) A speed that provides the maneuvering capability specified in 
Sec. 25.143(h) with the landing ice accretion defined in appendix C.
    (3) Changes in configuration, power or thrust, and speed, must be 
made in accordance with the established procedures for service 
operation.
    (4) The landing must be made without excessive vertical 
acceleration, tendency to bounce, nose over, ground loop, porpoise, or 
water loop.
    (5) The landings may not require exceptional piloting skill or 
alertness.
    (c) For landplanes and amphibians, the landing distance on land must 
be determined on a level, smooth, dry, hard-surfaced runway. In 
addition--
    (1) The pressures on the wheel braking systems may not exceed those 
specified by the brake manufacturer;
    (2) The brakes may not be used so as to cause excessive wear of 
brakes or tires; and
    (3) Means other than wheel brakes may be used if that means--
    (i) Is safe and reliable;
    (ii) Is used so that consistent results can be expected in service; 
and
    (iii) Is such that exceptional skill is not required to control the 
airplane.
    (d) For seaplanes and amphibians, the landing distance on water must 
be determined on smooth water.
    (e) For skiplanes, the landing distance on snow must be determined 
on smooth, dry, snow.
    (f) The landing distance data must include correction factors for 
not more than 50 percent of the nominal wind components along the 
landing path opposite to the direction of landing, and not less than 150 
percent of the nominal wind components along the landing path in the 
direction of landing.
    (g) If any device is used that depends on the operation of any 
engine, and if the landing distance would be noticeably increased when a 
landing is made with that engine inoperative, the landing distance must 
be determined with that engine inoperative unless the use of 
compensating means will result in a landing distance not more than that 
with each engine operating.

[Amdt. 25-121, 72 FR 44666; Aug. 8, 2007; 72 FR 50467, Aug. 31, 2007]

    Effective Date Note: At 79 FR 65525, Nov. 4, 2014, Sec. 25.125 was 
amended by revising paragraphs (a)(2), (b)(2)(ii)(B), and (b)(2)(ii)(C), 
effective Jan. 5, 2015. For the convenience of the user, the revised 
text is set forth as follows:



Sec. 25.125  Landing.

    (a) * * *
    (2) In icing conditions with the most critical of the landing ice 
accretion(s) defined in Appendices C and O of this part, as applicable, 
in accordance with Sec. 25.21(g), if VREF for icing 
conditions exceeds VREF for non-icing conditions by more than 
5 knots CAS at the maximum landing weight.
    (b) * * *
    (2) * * *
    (ii) * * *
    (B) 1.23 VSR0 with the most critical of the landing ice 
accretion(s) defined in Appendices C and O of this part, as applicable, 
in accordance with Sec. 25.21(g), if that speed exceeds VREF 
selected for non-icing conditions by more than 5 knots CAS; and
    (C) A speed that provides the maneuvering capability specified in 
Sec. 25.143(h) with the most critical of the landing ice accretion(s) 
defined in Appendices C and O of this part, as applicable, in accordance 
with Sec. 25.21(g).

                                * * * * *

                   Controllability and Maneuverability



Sec. 25.143  General.

    (a) The airplane must be safely controllable and maneuverable 
during--
    (1) Takeoff;
    (2) Climb;
    (3) Level flight;
    (4) Descent; and
    (5) Landing.
    (b) It must be possible to make a smooth transition from one flight 
condition to any other flight condition without exceptional piloting 
skill, alertness, or strength, and without danger of exceeding the 
airplane limit-load factor under any probable operating conditions, 
including--
    (1) The sudden failure of the critical engine;
    (2) For airplanes with three or more engines, the sudden failure of 
the second critical engine when the airplane is in the en route, 
approach, or landing

[[Page 386]]

configuration and is trimmed with the critical engine inoperative; and
    (3) Configuration changes, including deployment or retraction of 
deceleration devices.
    (c) The airplane must be shown to be safely controllable and 
maneuverable with the critical ice accretion appropriate to the phase of 
flight defined in appendix C, and with the critical engine inoperative 
and its propeller (if applicable) in the minimum drag position:
    (1) At the minimum V2 for takeoff;
    (2) During an approach and go-around; and
    (3) During an approach and landing.
    (d) The following table prescribes, for conventional wheel type 
controls, the maximum control forces permitted during the testing 
required by paragraph (a) through (c) of this section:

------------------------------------------------------------------------
   Force, in pounds, applied to the control
            wheel or rudder pedals              Pitch     Roll     Yaw
------------------------------------------------------------------------
For short term application for pitch and roll       75       50  .......
 control--two hands available for control....
For short term application for pitch and roll       50       25  .......
 control--one hand available for control.....
For short term application for yaw control...  .......  .......      150
For long term application....................       10        5       20
------------------------------------------------------------------------

    (e) Approved operating procedures or conventional operating 
practices must be followed when demonstrating compliance with the 
control force limitations for short term application that are prescribed 
in paragraph (d) of this section. The airplane must be in trim, or as 
near to being in trim as practical, in the preceding steady flight 
condition. For the takeoff condition, the airplane must be trimmed 
according to the approved operating procedures.
    (f) When demonstrating compliance with the control force limitations 
for long term application that are prescribed in paragraph (d) of this 
section, the airplane must be in trim, or as near to being in trim as 
practical.
    (g) When maneuvering at a constant airspeed or Mach number (up to 
VFC/MFC), the stick forces and the gradient of the 
stick force versus maneuvering load factor must lie within satisfactory 
limits. The stick forces must not be so great as to make excessive 
demands on the pilot's strength when maneuvering the airplane, and must 
not be so low that the airplane can easily be overstressed 
inadvertently. Changes of gradient that occur with changes of load 
factor must not cause undue difficulty in maintaining control of the 
airplane, and local gradients must not be so low as to result in a 
danger of overcontrolling.
    (h) The maneuvering capabilities in a constant speed coordinated 
turn at forward center of gravity, as specified in the following table, 
must be free of stall warning or other characteristics that might 
interfere with normal maneuvering:

--------------------------------------------------------------------------------------------------------------------------------------------------------
                                                                   Maneuvering bank
           Configuration                         Speed                angle in a                             Thrust/power setting
                                                                   coordinated turn
--------------------------------------------------------------------------------------------------------------------------------------------------------
Takeoff............................  V2                                    30[deg]   Asymmetric WAT-Limited. \1\
Takeoff............................  \2\V2 + XX                            40[deg]   All-engines-operating climb. \3\
En route...........................  VFTO                                  40[deg]   Asymmetric WAT-Limited. \1\
Landing............................  VREF                                  40[deg]   Symmetric for -3[deg] flight path angle.
--------------------------------------------------------------------------------------------------------------------------------------------------------
\1\ A combination of weight, altitude, and temperature (WAT) such that the thrust or power setting produces the minimum climb gradient specified in Sec.
    25.121 for the flight condition.
\2\ Airspeed approved for all-engines-operating initial climb.
\3\ That thrust or power setting which, in the event of failure of the critical engine and without any crew action to adjust the thrust or power of the
  remaining engines, would result in the thrust or power specified for the takeoff condition at V2, or any lesser thrust or power setting that is used
  for all-engines-operating initial climb procedures.

    (i) When demonstrating compliance with Sec. 25.143 in icing 
conditions--
    (1) Controllability must be demonstrated with the ice accretion 
defined in appendix C that is most critical for the particular flight 
phase;
    (2) It must be shown that a push force is required throughout a 
pushover maneuver down to a zero g load factor, or the lowest load 
factor obtainable if limited by elevator power or other design 
characteristic of the flight control system. It must be possible to 
promptly recover from the maneuver without exceeding a pull control 
force of 50 pounds; and

[[Page 387]]

    (3) Any changes in force that the pilot must apply to the pitch 
control to maintain speed with increasing sideslip angle must be 
steadily increasing with no force reversals, unless the change in 
control force is gradual and easily controllable by the pilot without 
using exceptional piloting skill, alertness, or strength.
    (j) For flight in icing conditions before the ice protection system 
has been activated and is performing its intended function, it must be 
demonstrated in flight with the ice accretion defined in appendix C, 
part II(e) of this part that:
    (1) The airplane is controllable in a pull-up maneuver up to 1.5 g 
load factor; and
    (2) There is no pitch control force reversal during a pushover 
maneuver down to 0.5 g load factor.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 
43 FR 2321, Jan. 16, 1978; Amdt. 25-84, 60 FR 30749, June 9, 1995; Amdt. 
25-108, 67 FR 70826, Nov. 26, 2002; Amdt. 25-121, 72 FR 44667, Aug. 8, 
2007; Amdt. 25-129, 74 FR 38339, Aug. 3, 2009]

    Effective Date Note: At 79 FR 65525, Nov. 4, 2014, Sec. 25.143 was 
amended by revising paragraphs (c) introductory text, (i)(1), and (j) 
introductory text, effective Jan. 5, 2015. For the convenience of the 
user, the revised text is set forth as follows:



Sec. 25.143  General.

                                * * * * *

    (c) The airplane must be shown to be safely controllable and 
maneuverable with the most critical of the ice accretion(s) appropriate 
to the phase of flight as defined in Appendices C and O of this part, as 
applicable, in accordance with Sec. 25.21(g), and with the critical 
engine inoperative and its propeller (if applicable) in the minimum drag 
position:

                                * * * * *

    (i) * * *
    (1) Controllability must be demonstrated with the most critical of 
the ice accretion(s) for the particular flight phase as defined in 
Appendices C and O of this part, as applicable, in accordance with Sec. 
25.21(g);

                                * * * * *

    (j) For flight in icing conditions before the ice protection system 
has been activated and is performing its intended function, it must be 
demonstrated in flight with the most critical of the ice accretion(s) 
defined in Appendix C, part II, paragraph (e) of this part and Appendix 
O, part II, paragraph (d) of this part, as applicable, in accordance 
with Sec. 25.21(g), that:

                                * * * * *



Sec. 25.145  Longitudinal control.

    (a) It must be possible, at any point between the trim speed 
prescribed in Sec. 25.103(b)(6) and stall identification (as defined in 
Sec. 25.201(d)), to pitch the nose downward so that the acceleration to 
this selected trim speed is prompt with
    (1) The airplane trimmed at the trim speed prescribed in Sec. 
25.103(b)(6);
    (2) The landing gear extended;
    (3) The wing flaps (i) retracted and (ii) extended; and
    (4) Power (i) off and (ii) at maximum continuous power on the 
engines.
    (b) With the landing gear extended, no change in trim control, or 
exertion of more than 50 pounds control force (representative of the 
maximum short term force that can be applied readily by one hand) may be 
required for the following maneuvers:
    (1) With power off, flaps retracted, and the airplane trimmed at 1.3 
VSR1, extend the flaps as rapidly as possible while 
maintaining the airspeed at approximately 30 percent above the reference 
stall speed existing at each instant throughout the maneuver.
    (2) Repeat paragraph (b)(1) except initially extend the flaps and 
then retract them as rapidly as possible.
    (3) Repeat paragraph (b)(2), except at the go-around power or thrust 
setting.
    (4) With power off, flaps retracted, and the airplane trimmed at 1.3 
VSR1, rapidly set go-around power or thrust while maintaining 
the same airspeed.
    (5) Repeat paragraph (b)(4) except with flaps extended.
    (6) With power off, flaps extended, and the airplane trimmed at 1.3 
VSR1, obtain and maintain airspeeds between VSW and either 1.6 VSR1 or 
VFE, whichever is lower.
    (c) It must be possible, without exceptional piloting skill, to 
prevent loss of altitude when complete retraction of the high lift 
devices from any position is begun during steady, straight, level flight 
at 1.08 VSR1 for propeller powered airplanes, or 1.13 VSR1 for turbojet 
powered airplanes, with--

[[Page 388]]

    (1) Simultaneous movement of the power or thrust controls to the go-
around power or thrust setting;
    (2) The landing gear extended; and
    (3) The critical combinations of landing weights and altitudes.
    (d) If gated high-lift device control positions are provided, 
paragraph (c) of this section applies to retractions of the high-lift 
devices from any position from the maximum landing position to the first 
gated position, between gated positions, and from the last gated 
position to the fully retracted position. The requirements of paragraph 
(c) of this section also apply to retractions from each approved landing 
position to the control position(s) associated with the high-lift device 
configuration(s) used to establish the go-around procedure(s) from that 
landing position. In addition, the first gated control position from the 
maximum landing position must correspond with a configuration of the 
high-lift devices used to establish a go-around procedure from a landing 
configuration. Each gated control position must require a separate and 
distinct motion of the control to pass through the gated position and 
must have features to prevent inadvertent movement of the control 
through the gated position. It must only be possible to make this 
separate and distinct motion once the control has reached the gated 
position.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5671, Apr. 8, 1970; Amdt. 25-72, 55 FR 29774, July 20, 1990; Amdt. 
25-84, 60 FR 30749, June 9, 1995; Amdt. 25-98, 64 FR 6164, Feb. 8, 1999; 
64 FR 10740, Mar. 5, 1999; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002]



Sec. 25.147  Directional and lateral control.

    (a) Directional control; general. It must be possible, with the 
wings level, to yaw into the operative engine and to safely make a 
reasonably sudden change in heading of up to 15 degrees in the direction 
of the critical inoperative engine. This must be shown at 1.3 VSR1 for 
heading changes up to 15 degrees (except that the heading change at 
which the rudder pedal force is 150 pounds need not be exceeded), and 
with--
    (1) The critical engine inoperative and its propeller in the minimum 
drag position;
    (2) The power required for level flight at 1.3 VSR1, but not more 
than maximum continuous power;
    (3) The most unfavorable center of gravity;
    (4) Landing gear retracted;
    (5) Flaps in the approach position; and
    (6) Maximum landing weight.
    (b) Directional control; airplanes with four or more engines. 
Airplanes with four or more engines must meet the requirements of 
paragraph (a) of this section except that--
    (1) The two critical engines must be inoperative with their 
propellers (if applicable) in the minimum drag position;
    (2) [Reserved]
    (3) The flaps must be in the most favorable climb position.
    (c) Lateral control; general. It must be possible to make 20[deg] 
banked turns, with and against the inoperative engine, from steady 
flight at a speed equal to 1.3 VSR1, with--
    (1) The critical engine inoperative and its propeller (if 
applicable) in the minimum drag position;
    (2) The remaining engines at maximum continuous power;
    (3) The most unfavorable center of gravity;
    (4) Landing gear (i) retracted and (ii) extended;
    (5) Flaps in the most favorable climb position; and
    (6) Maximum takeoff weight.
    (d) Lateral control; roll capability. With the critical engine 
inoperative, roll response must allow normal maneuvers. Lateral control 
must be sufficient, at the speeds likely to be used with one engine 
inoperative, to provide a roll rate necessary for safety without 
excessive control forces or travel.
    (e) Lateral control; airplanes with four or more engines. Airplanes 
with four or more engines must be able to make 20[deg] banked turns, 
with and against the inoperative engines, from steady flight at a speed 
equal to 1.3 VSR1, with maximum continuous power, and with the airplane 
in the configuration prescribed by paragraph (b) of this section.

[[Page 389]]

    (f) Lateral control; all engines operating. With the engines 
operating, roll response must allow normal maneuvers (such as recovery 
from upsets produced by gusts and the initiation of evasive maneuvers). 
There must be enough excess lateral control in sideslips (up to sideslip 
angles that might be required in normal operation), to allow a limited 
amount of maneuvering and to correct for gusts. Lateral control must be 
enough at any speed up to VFC/MFC to provide a peak roll rate necessary 
for safety, without excessive control forces or travel.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 
43 FR 2321, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990; 
Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-115, 69 FR 40527, 
July 2, 2004]



Sec. 25.149  Minimum control speed.

    (a) In establishing the minimum control speeds required by this 
section, the method used to simulate critical engine failure must 
represent the most critical mode of powerplant failure with respect to 
controllability expected in service.
    (b) VMC is the calibrated airspeed at which, when the critical 
engine is suddenly made inoperative, it is possible to maintain control 
of the airplane with that engine still inoperative and maintain straight 
flight with an angle of bank of not more than 5 degrees.
    (c) VMC may not exceed 1.13 VSR with--
    (1) Maximum available takeoff power or thrust on the engines;
    (2) The most unfavorable center of gravity;
    (3) The airplane trimmed for takeoff;
    (4) The maximum sea level takeoff weight (or any lesser weight 
necessary to show VMC);
    (5) The airplane in the most critical takeoff configuration existing 
along the flight path after the airplane becomes airborne, except with 
the landing gear retracted;
    (6) The airplane airborne and the ground effect negligible; and
    (7) If applicable, the propeller of the inoperative engine--
    (i) Windmilling;
    (ii) In the most probable position for the specific design of the 
propeller control; or
    (iii) Feathered, if the airplane has an automatic feathering device 
acceptable for showing compliance with the climb requirements of Sec. 
25.121.
    (d) The rudder forces required to maintain control at VMC may not 
exceed 150 pounds nor may it be necessary to reduce power or thrust of 
the operative engines. During recovery, the airplane may not assume any 
dangerous attitude or require exceptional piloting skill, alertness, or 
strength to prevent a heading change of more than 20 degrees.
    (e) VMCG, the minimum control speed on the ground, is the 
calibrated airspeed during the takeoff run at which, when the critical 
engine is suddenly made inoperative, it is possible to maintain control 
of the airplane using the rudder control alone (without the use of 
nosewheel steering), as limited by 150 pounds of force, and the lateral 
control to the extent of keeping the wings level to enable the takeoff 
to be safely continued using normal piloting skill. In the determination 
of VMCG, assuming that the path of the airplane accelerating with all 
engines operating is along the centerline of the runway, its path from 
the point at which the critical engine is made inoperative to the point 
at which recovery to a direction parallel to the centerline is completed 
may not deviate more than 30 feet laterally from the centerline at any 
point. VMCG must be established with--
    (1) The airplane in each takeoff configuration or, at the option of 
the applicant, in the most critical takeoff configuration;
    (2) Maximum available takeoff power or thrust on the operating 
engines;
    (3) The most unfavorable center of gravity;
    (4) The airplane trimmed for takeoff; and
    (5) The most unfavorable weight in the range of takeoff weights.
    (f) VMCL, the minimum control speed during approach and landing with 
all engines operating, is the calibrated airspeed at which, when the 
critical engine is suddenly made inoperative, it is possible to maintain 
control of the airplane with that engine still inoperative, and maintain 
straight flight with an angle of bank of not more than

[[Page 390]]

5 degrees. VMCL must be established with--
    (1) The airplane in the most critical configuration (or, at the 
option of the applicant, each configuration) for approach and landing 
with all engines operating;
    (2) The most unfavorable center of gravity;
    (3) The airplane trimmed for approach with all engines operating;
    (4) The most favorable weight, or, at the option of the applicant, 
as a function of weight;
    (5) For propeller airplanes, the propeller of the inoperative engine 
in the position it achieves without pilot action, assuming the engine 
fails while at the power or thrust necessary to maintain a three degree 
approach path angle; and
    (6) Go-around power or thrust setting on the operating engine(s).
    (g) For airplanes with three or more engines, VMCL-2, the minimum 
control speed during approach and landing with one critical engine 
inoperative, is the calibrated airspeed at which, when a second critical 
engine is suddenly made inoperative, it is possible to maintain control 
of the airplane with both engines still inoperative, and maintain 
straight flight with an angle of bank of not more than 5 degrees. VMCL-2 
must be established with--
    (1) The airplane in the most critical configuration (or, at the 
option of the applicant, each configuration) for approach and landing 
with one critical engine inoperative;
    (2) The most unfavorable center of gravity;
    (3) The airplane trimmed for approach with one critical engine 
inoperative;
    (4) The most unfavorable weight, or, at the option of the applicant, 
as a function of weight;
    (5) For propeller airplanes, the propeller of the more critical 
inoperative engine in the position it achieves without pilot action, 
assuming the engine fails while at the power or thrust necessary to 
maintain a three degree approach path angle, and the propeller of the 
other inoperative engine feathered;
    (6) The power or thrust on the operating engine(s) necessary to 
maintain an approach path angle of three degrees when one critical 
engine is inoperative; and
    (7) The power or thrust on the operating engine(s) rapidly changed, 
immediately after the second critical engine is made inoperative, from 
the power or thrust prescribed in paragraph (g)(6) of this section to--
    (i) Minimum power or thrust; and
    (ii) Go-around power or thrust setting.
    (h) In demonstrations of VMCL and VMCL-2--
    (1) The rudder force may not exceed 150 pounds;
    (2) The airplane may not exhibit hazardous flight characteristics or 
require exceptional piloting skill, alertness, or strength;
    (3) Lateral control must be sufficient to roll the airplane, from an 
initial condition of steady flight, through an angle of 20 degrees in 
the direction necessary to initiate a turn away from the inoperative 
engine(s), in not more than 5 seconds; and
    (4) For propeller airplanes, hazardous flight characteristics must 
not be exhibited due to any propeller position achieved when the engine 
fails or during any likely subsequent movements of the engine or 
propeller controls.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 
43 FR 2321, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990; 55 
FR 37607, Sept. 12, 1990; Amdt. 25-84, 60 FR 30749, June 9, 1995; Amdt. 
25-108, 67 FR 70827, Nov. 26, 2002]

                                  Trim



Sec. 25.161  Trim.

    (a) General. Each airplane must meet the trim requirements of this 
section after being trimmed, and without further pressure upon, or 
movement of, either the primary controls or their corresponding trim 
controls by the pilot or the automatic pilot.
    (b) Lateral and directional trim. The airplane must maintain lateral 
and directional trim with the most adverse lateral displacement of the 
center of gravity within the relevant operating limitations, during 
normally expected conditions of operation (including operation at any 
speed from 1.3 VSR1 to VMO/MMO).

[[Page 391]]

    (c) Longitudinal trim. The airplane must maintain longitudinal trim 
during--
    (1) A climb with maximum continuous power at a speed not more than 
1.3 VSR1, with the landing gear retracted, and the flaps (i) 
retracted and (ii) in the takeoff position;
    (2) Either a glide with power off at a speed not more than 1.3 
VSR1, or an approach within the normal range of approach 
speeds appropriate to the weight and configuration with power settings 
corresponding to a 3 degree glidepath, whichever is the most severe, 
with the landing gear extended, the wing flaps (i) retracted and (ii) 
extended, and with the most unfavorable combination of center of gravity 
position and weight approved for landing; and
    (3) Level flight at any speed from 1.3 VSR1, to VMO/MMO, 
with the landing gear and flaps retracted, and from 1.3 VSR1 
to VLE with the landing gear extended.
    (d) Longitudinal, directional, and lateral trim. The airplane must 
maintain longitudinal, directional, and lateral trim (and for the 
lateral trim, the angle of bank may not exceed five degrees) at 1.3 
VSR1 during climbing flight with--
    (1) The critical engine inoperative;
    (2) The remaining engines at maximum continuous power; and
    (3) The landing gear and flaps retracted.
    (e) Airplanes with four or more engines. Each airplane with four or 
more engines must also maintain trim in rectilinear flight with the most 
unfavorable center of gravity and at the climb speed, configuration, and 
power required by Sec. 25.123(a) for the purpose of establishing the en 
route flight paths with two engines inoperative.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5671, Apr. 8, 1970; Amdt. 25-38, 41 FR 55466, Dec. 20, 1976; Amdt. 
25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-115, 69 FR 40527, July 2, 
2004]

                                Stability



Sec. 25.171  General.

    The airplane must be longitudinally, directionally, and laterally 
stable in accordance with the provisions of Sec. Sec. 25.173 through 
25.177. In addition, suitable stability and control feel (static 
stability) is required in any condition normally encountered in service, 
if flight tests show it is necessary for safe operation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 
FR 13117, Oct. 15, 1965]



Sec. 25.173  Static longitudinal stability.

    Under the conditions specified in Sec. 25.175, the characteristics 
of the elevator control forces (including friction) must be as follows:
    (a) A pull must be required to obtain and maintain speeds below the 
specified trim speed, and a push must be required to obtain and maintain 
speeds above the specified trim speed. This must be shown at any speed 
that can be obtained except speeds higher than the landing gear or wing 
flap operating limit speeds or VFC/MFC, whichever is appropriate, or 
lower than the minimum speed for steady unstalled flight.
    (b) The airspeed must return to within 10 percent of the original 
trim speed for the climb, approach, and landing conditions specified in 
Sec. 25.175 (a), (c), and (d), and must return to within 7.5 percent of 
the original trim speed for the cruising condition specified in Sec. 
25.175(b), when the control force is slowly released from any speed 
within the range specified in paragraph (a) of this section.
    (c) The average gradient of the stable slope of the stick force 
versus speed curve may not be less than 1 pound for each 6 knots.
    (d) Within the free return speed range specified in paragraph (b) of 
this section, it is permissible for the airplane, without control 
forces, to stabilize on speeds above or below the desired trim speeds if 
exceptional attention on the part of the pilot is not required to return 
to and maintain the desired trim speed and altitude.

[Amdt. 25-7, 30 FR 13117, Oct. 15, 1965]



Sec. 25.175  Demonstration of static longitudinal stability.

    Static longitudinal stability must be shown as follows:
    (a) Climb. The stick force curve must have a stable slope at speeds 
between

[[Page 392]]

85 and 115 percent of the speed at which the airplane--
    (1) Is trimmed, with--
    (i) Wing flaps retracted;
    (ii) Landing gear retracted;
    (iii) Maximum takeoff weight; and
    (iv) 75 percent of maximum continuous power for reciprocating 
engines or the maximum power or thrust selected by the applicant as an 
operating limitation for use during climb for turbine engines; and
    (2) Is trimmed at the speed for best rate-of-climb except that the 
speed need not be less than 1.3 VSR1.
    (b) Cruise. Static longitudinal stability must be shown in the 
cruise condition as follows:
    (1) With the landing gear retracted at high speed, the stick force 
curve must have a stable slope at all speeds within a range which is the 
greater of 15 percent of the trim speed plus the resulting free return 
speed range, or 50 knots plus the resulting free return speed range, 
above and below the trim speed (except that the speed range need not 
include speeds less than 1.3 VSR1, nor speeds greater than 
VFC/MFC, nor speeds that require a stick force of more than 50 pounds), 
with--
    (i) The wing flaps retracted;
    (ii) The center of gravity in the most adverse position (see Sec. 
25.27);
    (iii) The most critical weight between the maximum takeoff and 
maximum landing weights;
    (iv) 75 percent of maximum continuous power for reciprocating 
engines or for turbine engines, the maximum cruising power selected by 
the applicant as an operating limitation (see Sec. 25.1521), except 
that the power need not exceed that required at VMO/MMO; and
    (v) The airplane trimmed for level flight with the power required in 
paragraph (b)(1)(iv) of this section.
    (2) With the landing gear retracted at low speed, the stick force 
curve must have a stable slope at all speeds within a range which is the 
greater of 15 percent of the trim speed plus the resulting free return 
speed range, or 50 knots plus the resulting free return speed range, 
above and below the trim speed (except that the speed range need not 
include speeds less than 1.3 VSR1, nor speeds greater than 
the minimum speed of the applicable speed range prescribed in paragraph 
(b)(1), nor speeds that require a stick force of more than 50 pounds), 
with--
    (i) Wing flaps, center of gravity position, and weight as specified 
in paragraph (b)(1) of this section;
    (ii) Power required for level flight at a speed equal to (VMO + 1.3 
VSR1)/2; and
    (iii) The airplane trimmed for level flight with the power required 
in paragraph (b)(2)(ii) of this section.
    (3) With the landing gear extended, the stick force curve must have 
a stable slope at all speeds within a range which is the greater of 15 
percent of the trim speed plus the resulting free return speed range, or 
50 knots plus the resulting free return speed range, above and below the 
trim speed (except that the speed range need not include speeds less 
than 1.3 VSR1, nor speeds greater than VLE, nor speeds that 
require a stick force of more than 50 pounds), with--
    (i) Wing flap, center of gravity position, and weight as specified 
in paragraph (b)(1) of this section;
    (ii) 75 percent of maximum continuous power for reciprocating 
engines or, for turbine engines, the maximum cruising power selected by 
the applicant as an operating limitation, except that the power need not 
exceed that required for level flight at VLE; and
    (iii) The aircraft trimmed for level flight with the power required 
in paragraph (b)(3)(ii) of this section.
    (c) Approach. The stick force curve must have a stable slope at 
speeds between VSW and 1.7 VSR1, with--
    (1) Wing flaps in the approach position;
    (2) Landing gear retracted;
    (3) Maximum landing weight; and
    (4) The airplane trimmed at 1.3 VSR1 with enough power to 
maintain level flight at this speed.
    (d) Landing. The stick force curve must have a stable slope, and the 
stick force may not exceed 80 pounds, at speeds between VSW and 1.7 
VSR0 with--
    (1) Wing flaps in the landing position;
    (2) Landing gear extended;
    (3) Maximum landing weight;
    (4) The airplane trimmed at 1.3 VSR0 with--
    (i) Power or thrust off, and
    (ii) Power or thrust for level flight.

[[Page 393]]

    (5) The airplane trimmed at 1.3 VSR0 with power or thrust 
off.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 
FR 13117, Oct. 15, 1965; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 
25-115, 69 FR 40527, July 2, 2004]



Sec. 25.177  Static lateral-directional stability.

    (a) The static directional stability (as shown by the tendency to 
recover from a skid with the rudder free) must be positive for any 
landing gear and flap position and symmetric power condition, at speeds 
from 1.13 VSR1, up to VFE, VLE, or 
VFC/MFC (as appropriate for the airplane 
configuration).
    (b) The static lateral stability (as shown by the tendency to raise 
the low wing in a sideslip with the aileron controls free) for any 
landing gear and flap position and symmetric power condition, may not be 
negative at any airspeed (except that speeds higher than VFE 
need not be considered for flaps extended configurations nor speeds 
higher than VLE for landing gear extended configurations) in 
the following airspeed ranges:
    (1) From 1.13 VSR1 to VMO/MMO.
    (2) From VMO/MMO to VFC/
MFC, unless the divergence is--
    (i) Gradual;
    (ii) Easily recognizable by the pilot; and
    (iii) Easily controllable by the pilot.
    (c) The following requirement must be met for the configurations and 
speed specified in paragraph (a) of this section. In straight, steady 
sideslips over the range of sideslip angles appropriate to the operation 
of the airplane, the aileron and rudder control movements and forces 
must be substantially proportional to the angle of sideslip in a stable 
sense. This factor of proportionality must lie between limits found 
necessary for safe operation. The range of sideslip angles evaluated 
must include those sideslip angles resulting from the lesser of:
    (1) One-half of the available rudder control input; and
    (2) A rudder control force of 180 pounds.
    (d) For sideslip angles greater than those prescribed by paragraph 
(c) of this section, up to the angle at which full rudder control is 
used or a rudder control force of 180 pounds is obtained, the rudder 
control forces may not reverse, and increased rudder deflection must be 
needed for increased angles of sideslip. Compliance with this 
requirement must be shown using straight, steady sideslips, unless full 
lateral control input is achieved before reaching either full rudder 
control input or a rudder control force of 180 pounds; a straight, 
steady sideslip need not be maintained after achieving full lateral 
control input. This requirement must be met at all approved landing gear 
and flap positions for the range of operating speeds and power 
conditions appropriate to each landing gear and flap position with all 
engines operating.

[Amdt. 25-135, 76 FR 74654, Dec. 1, 2011]



Sec. 25.181  Dynamic stability.

    (a) Any short period oscillation, not including combined lateral-
directional oscillations, occurring between 1.13 VSR and 
maximum allowable speed appropriate to the configuration of the airplane 
must be heavily damped with the primary controls--
    (1) Free; and
    (2) In a fixed position.
    (b) Any combined lateral-directional oscillations (``Dutch roll'') 
occurring between 1.13 VSR and maximum allowable speed 
appropriate to the configuration of the airplane must be positively 
damped with controls free, and must be controllable with normal use of 
the primary controls without requiring exceptional pilot skill.

[Amdt. 25-42, 43 FR 2322, Jan. 16, 1978, as amended by Amdt. 25-72, 55 
FR 29775, July 20, 1990; 55 FR 37607, Sept. 12, 1990; Amdt. 25-108, 67 
FR 70827, Nov. 26, 2002]

                                 Stalls



Sec. 25.201  Stall demonstration.

    (a) Stalls must be shown in straight flight and in 30 degree banked 
turns with--
    (1) Power off; and
    (2) The power necessary to maintain level flight at 1.5 VSR1 (where 
VSR1 corresponds to the reference stall speed at maximum landing weight 
with flaps in the approach position and the landing gear retracted).
    (b) In each condition required by paragraph (a) of this section, it 
must

[[Page 394]]

be possible to meet the applicable requirements of Sec. 25.203 with--
    (1) Flaps, landing gear, and deceleration devices in any likely 
combination of positions approved for operation;
    (2) Representative weights within the range for which certification 
is requested;
    (3) The most adverse center of gravity for recovery; and
    (4) The airplane trimmed for straight flight at the speed prescribed 
in Sec. 25.103(b)(6).
    (c) The following procedures must be used to show compliance with 
Sec. 25.203;
    (1) Starting at a speed sufficiently above the stalling speed to 
ensure that a steady rate of speed reduction can be established, apply 
the longitudinal control so that the speed reduction does not exceed one 
knot per second until the airplane is stalled.
    (2) In addition, for turning flight stalls, apply the longitudinal 
control to achieve airspeed deceleration rates up to 3 knots per second.
    (3) As soon as the airplane is stalled, recover by normal recovery 
techniques.
    (d) The airplane is considered stalled when the behavior of the 
airplane gives the pilot a clear and distinctive indication of an 
acceptable nature that the airplane is stalled. Acceptable indications 
of a stall, occurring either individually or in combination, are--
    (1) A nose-down pitch that cannot be readily arrested;
    (2) Buffeting, of a magnitude and severity that is a strong and 
effective deterrent to further speed reduction; or
    (3) The pitch control reaches the aft stop and no further increase 
in pitch attitude occurs when the control is held full aft for a short 
time before recovery is initiated.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-84, 
60 FR 30750, June 9, 1995; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002]



Sec. 25.203  Stall characteristics.

    (a) It must be possible to produce and to correct roll and yaw by 
unreversed use of the aileron and rudder controls, up to the time the 
airplane is stalled. No abnormal nose-up pitching may occur. The 
longitudinal control force must be positive up to and throughout the 
stall. In addition, it must be possible to promptly prevent stalling and 
to recover from a stall by normal use of the controls.
    (b) For level wing stalls, the roll occurring between the stall and 
the completion of the recovery may not exceed approximately 20 degrees.
    (c) For turning flight stalls, the action of the airplane after the 
stall may not be so violent or extreme as to make it difficult, with 
normal piloting skill, to effect a prompt recovery and to regain control 
of the airplane. The maximum bank angle that occurs during the recovery 
may not exceed--
    (1) Approximately 60 degrees in the original direction of the turn, 
or 30 degrees in the opposite direction, for deceleration rates up to 1 
knot per second; and
    (2) Approximately 90 degrees in the original direction of the turn, 
or 60 degrees in the opposite direction, for deceleration rates in 
excess of 1 knot per second.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-84, 
60 FR 30750, June 9, 1995]



Sec. 25.207  Stall warning.

    (a) Stall warning with sufficient margin to prevent inadvertent 
stalling with the flaps and landing gear in any normal position must be 
clear and distinctive to the pilot in straight and turning flight.
    (b) The warning must be furnished either through the inherent 
aerodynamic qualities of the airplane or by a device that will give 
clearly distinguishable indications under expected conditions of flight. 
However, a visual stall warning device that requires the attention of 
the crew within the cockpit is not acceptable by itself. If a warning 
device is used, it must provide a warning in each of the airplane 
configurations prescribed in paragraph (a) of this section at the speed 
prescribed in paragraphs (c) and (d) of this section. Except for showing 
compliance with the stall warning margin prescribed in paragraph 
(h)(3)(ii) of this section, stall warning for flight in icing conditions 
must be provided by the same means as stall warning for flight in non-
icing conditions.
    (c) When the speed is reduced at rates not exceeding one knot per 
second,

[[Page 395]]

stall warning must begin, in each normal configuration, at a speed, 
VSW, exceeding the speed at which the stall is identified in 
accordance with Sec. 25.201(d) by not less than five knots or five 
percent CAS, whichever is greater. Once initiated, stall warning must 
continue until the angle of attack is reduced to approximately that at 
which stall warning began.
    (d) In addition to the requirement of paragraph (c) of this section, 
when the speed is reduced at rates not exceeding one knot per second, in 
straight flight with engines idling and at the center-of-gravity 
position specified in Sec. 25.103(b)(5), VSW, in each normal 
configuration, must exceed VSR by not less than three knots 
or three percent CAS, whichever is greater.
    (e) In icing conditions, the stall warning margin in straight and 
turning flight must be sufficient to allow the pilot to prevent stalling 
(as defined in Sec. 25.201(d)) when the pilot starts a recovery 
maneuver not less than three seconds after the onset of stall warning. 
When demonstrating compliance with this paragraph, the pilot must 
perform the recovery maneuver in the same way as for the airplane in 
non-icing conditions. Compliance with this requirement must be 
demonstrated in flight with the speed reduced at rates not exceeding one 
knot per second, with--
    (1) The more critical of the takeoff ice and final takeoff ice 
accretions defined in appendix C for each configuration used in the 
takeoff phase of flight;
    (2) The en route ice accretion defined in appendix C for the en 
route configuration;
    (3) The holding ice accretion defined in appendix C for the holding 
configuration(s);
    (4) The approach ice accretion defined in appendix C for the 
approach configuration(s); and
    (5) The landing ice accretion defined in appendix C for the landing 
and go-around configuration(s).
    (f) The stall warning margin must be sufficient in both non-icing 
and icing conditions to allow the pilot to prevent stalling when the 
pilot starts a recovery maneuver not less than one second after the 
onset of stall warning in slow-down turns with at least 1.5 g load 
factor normal to the flight path and airspeed deceleration rates of at 
least 2 knots per second. When demonstrating compliance with this 
paragraph for icing conditions, the pilot must perform the recovery 
maneuver in the same way as for the airplane in non-icing conditions. 
Compliance with this requirement must be demonstrated in flight with--
    (1) The flaps and landing gear in any normal position;
    (2) The airplane trimmed for straight flight at a speed of 1.3 
VSR; and
    (3) The power or thrust necessary to maintain level flight at 1.3 
VSR.
    (g) Stall warning must also be provided in each abnormal 
configuration of the high lift devices that is likely to be used in 
flight following system failures (including all configurations covered 
by Airplane Flight Manual procedures).
    (h) For flight in icing conditions before the ice protection system 
has been activated and is performing its intended function, with the ice 
accretion defined in appendix C, part II(e) of this part, the stall 
warning margin in straight and turning flight must be sufficient to 
allow the pilot to prevent stalling without encountering any adverse 
flight characteristics when:
    (1) The speed is reduced at rates not exceeding one knot per second;
    (2) The pilot performs the recovery maneuver in the same way as for 
flight in non-icing conditions; and
    (3) The recovery maneuver is started no earlier than:
    (i) One second after the onset of stall warning if stall warning is 
provided by the same means as for flight in non-icing conditions; or
    (ii) Three seconds after the onset of stall warning if stall warning 
is provided by a different means than for flight in non-icing 
conditions.
    (i) In showing compliance with paragraph (h) of this section, if 
stall warning is provided by a different means in icing conditions than 
for non-icing conditions, compliance with Sec. 25.203 must be shown 
using the accretion defined in appendix C, part II(e) of this part. 
Compliance with this requirement must be

[[Page 396]]

shown using the demonstration prescribed by Sec. 25.201, except that 
the deceleration rates of Sec. 25.201(c)(2) need not be demonstrated.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 
FR 13118, Oct. 15, 1965; Amdt. 25-42, 43 FR 2322, Jan. 16, 1978; Amdt. 
25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-121, 72 FR 44668, Aug. 8, 
2007; Amdt. 25-129, 74 FR 38339, Aug. 3, 2009]

    Effective Date Note: At 79 FR 65526, Nov. 4, 2014, Sec. 25.207 was 
amended by revising paragraphs (b), (e)(1), (e)(2), (e)(3), (e)(4), 
(e)(5), and (h) introductory text, effective Jan. 5, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.207  Stall warning.

                                * * * * *

    (b) The warning must be furnished either through the inherent 
aerodynamic qualities of the airplane or by a device that will give 
clearly distinguishable indications under expected conditions of flight. 
However, a visual stall warning device that requires the attention of 
the crew within the cockpit is not acceptable by itself. If a warning 
device is used, it must provide a warning in each of the airplane 
configurations prescribed in paragraph (a) of this section at the speed 
prescribed in paragraphs (c) and (d) of this section. Except for the 
stall warning prescribed in paragraph (h)(3)(ii) of this section, the 
stall warning for flight in icing conditions must be provided by the 
same means as the stall warning for flight in non-icing conditions.

                                * * * * *

    (e) * * *
    (1) The most critical of the takeoff ice and final takeoff ice 
accretions defined in Appendices C and O of this part, as applicable, in 
accordance with Sec. 25.21(g), for each configuration used in the 
takeoff phase of flight;
    (2) The most critical of the en route ice accretion(s) defined in 
Appendices C and O of this part, as applicable, in accordance with Sec. 
25.21(g), for the en route configuration;
    (3) The most critical of the holding ice accretion(s) defined in 
Appendices C and O of this part, as applicable, in accordance with Sec. 
25.21(g), for the holding configuration(s);
    (4) The most critical of the approach ice accretion(s) defined in 
Appendices C and O of this part, as applicable, in accordance with Sec. 
25.21(g), for the approach configuration(s); and
    (5) The most critical of the landing ice accretion(s) defined in 
Appendices C and O of this part, as applicable, in accordance with Sec. 
25.21(g), for the landing and go-around configuration(s).

                                * * * * *

    (h) The following stall warning margin is required for flight in 
icing conditions before the ice protection system has been activated and 
is performing its intended function. Compliance must be shown using the 
most critical of the ice accretion(s) defined in Appendix C, part II, 
paragraph (e) of this part and Appendix O, part II, paragraph (d) of 
this part, as applicable, in accordance with Sec. 25.21(g). The stall 
warning margin in straight and turning flight must be sufficient to 
allow the pilot to prevent stalling without encountering any adverse 
flight characteristics when:

                                * * * * *

                Ground and Water Handling Characteristics



Sec. 25.231  Longitudinal stability and control.

    (a) Landplanes may have no uncontrollable tendency to nose over in 
any reasonably expected operating condition or when rebound occurs 
during landing or takeoff. In addition--
    (1) Wheel brakes must operate smoothly and may not cause any undue 
tendency to nose over; and
    (2) If a tail-wheel landing gear is used, it must be possible, 
during the takeoff ground run on concrete, to maintain any attitude up 
to thrust line level, at 75 percent of VSR1.
    (b) For seaplanes and amphibians, the most adverse water conditions 
safe for takeoff, taxiing, and landing, must be established.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-108, 
67 FR 70828, Nov. 26, 2002]



Sec. 25.233  Directional stability and control.

    (a) There may be no uncontrollable ground-looping tendency in 
90[deg] cross winds, up to a wind velocity of 20 knots or 0.2 
VSR0, whichever is greater, except that the wind velocity 
need not exceed 25 knots at any speed at which the airplane may be 
expected to be operated on the ground. This may be shown while 
establishing the 90[deg] cross component of wind velocity required by 
Sec. 25.237.

[[Page 397]]

    (b) Landplanes must be satisfactorily controllable, without 
exceptional piloting skill or alertness, in power-off landings at normal 
landing speed, without using brakes or engine power to maintain a 
straight path. This may be shown during power-off landings made in 
conjunction with other tests.
    (c) The airplane must have adequate directional control during 
taxiing. This may be shown during taxiing prior to takeoffs made in 
conjunction with other tests.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5671, Apr. 8, 1970; Amdt. 25-42, 43 FR 2322, Jan. 16, 1978; Amdt. 
25-94, 63 FR 8848, Feb. 23, 1998; Amdt. 25-108, 67 FR 70828, Nov. 26, 
2002]



Sec. 25.235  Taxiing condition.

    The shock absorbing mechanism may not damage the structure of the 
airplane when the airplane is taxied on the roughest ground that may 
reasonably be expected in normal operation.



Sec. 25.237  Wind velocities.

    (a) For land planes and amphibians, the following applies:
    (1) A 90-degree cross component of wind velocity, demonstrated to be 
safe for takeoff and landing, must be established for dry runways and 
must be at least 20 knots or 0.2 VSR0, whichever is greater, 
except that it need not exceed 25 knots.
    (2) The crosswind component for takeoff established without ice 
accretions is valid in icing conditions.
    (3) The landing crosswind component must be established for:
    (i) Non-icing conditions, and
    (ii) Icing conditions with the landing ice accretion defined in 
appendix C.
    (b) For seaplanes and amphibians, the following applies:
    (1) A 90-degree cross component of wind velocity, up to which 
takeoff and landing is safe under all water conditions that may 
reasonably be expected in normal operation, must be established and must 
be at least 20 knots or 0.2 VSR0, whichever is greater, 
except that it need not exceed 25 knots.
    (2) A wind velocity, for which taxiing is safe in any direction 
under all water conditions that may reasonably be expected in normal 
operation, must be established and must be at least 20 knots or 0.2 
VSR0, whichever is greater, except that it need not exceed 25 knots.

[Amdt. 25-42, 43 FR 2322, Jan. 16, 1978, as amended by Amdt. 25-108, 67 
FR 70827, Nov. 26, 2002; Amdt. 25-121, 72 FR 44668, Aug. 8, 2007]

    Effective Date Note: At 79 FR 65526, Nov. 4, 2014, Sec. 25.237 was 
amended by revising paragraph (a)(3)(ii), effective Jan. 5, 2015. For 
the convenience of the user, the revised text is set forth as follows:



Sec. 25.237  Wind velocities.

    (a) * * *
    (3) * * *
    (ii) Icing conditions with the most critical of the landing ice 
accretion(s) defined in Appendices C and O of this part, as applicable, 
in accordance with Sec. 25.21(g).

                                * * * * *



Sec. 25.239  Spray characteristics, control, and stability on water.

    (a) For seaplanes and amphibians, during takeoff, taxiing, and 
landing, and in the conditions set forth in paragraph (b) of this 
section, there may be no--
    (1) Spray characteristics that would impair the pilot's view, cause 
damage, or result in the taking in of an undue quantity of water;
    (2) Dangerously uncontrollable porpoising, bounding, or swinging 
tendency; or
    (3) Immersion of auxiliary floats or sponsons, wing tips, propeller 
blades, or other parts not designed to withstand the resulting water 
loads.
    (b) Compliance with the requirements of paragraph (a) of this 
section must be shown--
    (1) In water conditions, from smooth to the most adverse condition 
established in accordance with Sec. 25.231;
    (2) In wind and cross-wind velocities, water currents, and 
associated waves and swells that may reasonably be expected in operation 
on water;
    (3) At speeds that may reasonably be expected in operation on water;
    (4) With sudden failure of the critical engine at any time while on 
water; and
    (5) At each weight and center of gravity position, relevant to each 
operating condition, within the range of loading conditions for which 
certification is requested.

[[Page 398]]

    (c) In the water conditions of paragraph (b) of this section, and in 
the corresponding wind conditions, the seaplane or amphibian must be 
able to drift for five minutes with engines inoperative, aided, if 
necessary, by a sea anchor.

                    Miscellaneous Flight Requirements



Sec. 25.251  Vibration and buffeting.

    (a) The airplane must be demonstrated in flight to be free from any 
vibration and buffeting that would prevent continued safe flight in any 
likely operating condition.
    (b) Each part of the airplane must be demonstrated in flight to be 
free from excessive vibration under any appropriate speed and power 
conditions up to VDF/MDF. The maximum speeds shown 
must be used in establishing the operating limitations of the airplane 
in accordance with Sec. 25.1505.
    (c) Except as provided in paragraph (d) of this section, there may 
be no buffeting condition, in normal flight, including configuration 
changes during cruise, severe enough to interfere with the control of 
the airplane, to cause excessive fatigue to the crew, or to cause 
structural damage. Stall warning buffeting within these limits is 
allowable.
    (d) There may be no perceptible buffeting condition in the cruise 
configuration in straight flight at any speed up to VMO/MMO, except that 
stall warning buffeting is allowable.
    (e) For an airplane with MD greater than .6 or with a 
maximum operating altitude greater than 25,000 feet, the positive 
maneuvering load factors at which the onset of perceptible buffeting 
occurs must be determined with the airplane in the cruise configuration 
for the ranges of airspeed or Mach number, weight, and altitude for 
which the airplane is to be certificated. The envelopes of load factor, 
speed, altitude, and weight must provide a sufficient range of speeds 
and load factors for normal operations. Probable inadvertent excursions 
beyond the boundaries of the buffet onset envelopes may not result in 
unsafe conditions.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5671, Apr. 8, 1970; Amdt. 25-72, 55 FR 29775, July 20, 1990; Amdt. 
25-77, 57 FR 28949, June 29, 1992]



Sec. 25.253  High-speed characteristics.

    (a) Speed increase and recovery characteristics. The following speed 
increase and recovery characteristics must be met:
    (1) Operating conditions and characteristics likely to cause 
inadvertent speed increases (including upsets in pitch and roll) must be 
simulated with the airplane trimmed at any likely cruise speed up to 
VMO/MMO. These conditions and characteristics include gust upsets, 
inadvertent control movements, low stick force gradient in relation to 
control friction, passenger movement, leveling off from climb, and 
descent from Mach to airspeed limit altitudes.
    (2) Allowing for pilot reaction time after effective inherent or 
artificial speed warning occurs, it must be shown that the airplane can 
be recovered to a normal attitude and its speed reduced to VMO/MMO, 
without--
    (i) Exceptional piloting strength or skill;
    (ii) Exceeding VD/MD, VDF/MDF, or the structural limitations; and
    (iii) Buffeting that would impair the pilot's ability to read the 
instruments or control the airplane for recovery.
    (3) With the airplane trimmed at any speed up to VMO/
MMO, there must be no reversal of the response to control 
input about any axis at any speed up to VDF/MDF. 
Any tendency to pitch, roll, or yaw must be mild and readily 
controllable, using normal piloting techniques. When the airplane is 
trimmed at VMO/MMO, the slope of the elevator 
control force versus speed curve need not be stable at speeds greater 
than VFC/MFC, but there must be a push force at 
all speeds up to VDF/MDF and there must be no 
sudden or excessive reduction of elevator control force as 
VDF/MDF is reached.
    (4) Adequate roll capability to assure a prompt recovery from a 
lateral upset condition must be available at any speed up to 
VDF/MDF.
    (5) With the airplane trimmed at VMO/MMO, 
extension of the speedbrakes over the available range of movements of 
the pilot's control, at all speeds above VMO/MMO, 
but not so high that VDF/MDF would be exceeded 
during the maneuver, must not result in:

[[Page 399]]

    (i) An excessive positive load factor when the pilot does not take 
action to counteract the effects of extension;
    (ii) Buffeting that would impair the pilot's ability to read the 
instruments or control the airplane for recovery; or
    (iii) A nose down pitching moment, unless it is small.
    (b) Maximum speed for stability characteristics, VFC/MFC. 
VFC/MFC is the maximum speed at which the 
requirements of Sec. Sec. 25.143(g), 25.147(f), 25.175(b)(1), 25.177(a) 
through (c), and 25.181 must be met with flaps and landing gear 
retracted. Except as noted in Sec. 25.253(c), VFC/
MFC may not be less than a speed midway between 
VMO/MMO and VDF/MDF, except 
that, for altitudes where Mach number is the limiting factor, 
MFC need not exceed the Mach number at which effective speed 
warning occurs.
    (c) Maximum speed for stability characteristics in icing conditions. 
The maximum speed for stability characteristics with the ice accretions 
defined in appendix C, at which the requirements of Sec. Sec. 
25.143(g), 25.147(f), 25.175(b)(1), 25.177(a) through (c), and 25.181 
must be met, is the lower of:
    (1) 300 knots CAS;
    (2) VFC; or
    (3) A speed at which it is demonstrated that the airframe will be 
free of ice accretion due to the effects of increased dynamic pressure.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5671, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; 
Amdt. 25-72, 55 FR 29775, July 20, 1990; Amdt. 25-84, 60 FR 30750, June 
9, 1995; Amdt. 25-121, 72 FR 44668, Aug. 8, 2007; Amdt. 25-135, 76 FR 
74654, Dec. 1, 2011]

    Effective Date Note: At 79 FR 65526, Nov. 4, 2014, Sec. 25.253 was 
amended by revising paragraph (c) introductory text, effective Jan. 5. 
2015. For the convenience of the user, the revised text is set forth as 
follows:



Sec. 25.253  High-speed characteristics.

                                * * * * *

    (c) Maximum speed for stability characteristics in icing conditions. 
The maximum speed for stability characteristics with the most critical 
of the ice accretions defined in Appendices C and O of this part, as 
applicable, in accordance with Sec. 25.21(g), at which the requirements 
of Sec. Sec. 25.143(g), 25.147(f), 25.175(b)(1), 25.177(a) through (c), 
and 25.181 must be met, is the lower of:

                                * * * * *



Sec. 25.255  Out-of-trim characteristics.

    (a) From an initial condition with the airplane trimmed at cruise 
speeds up to VMO/MMO, the airplane must have satisfactory maneuvering 
stability and controllability with the degree of out-of-trim in both the 
airplane nose-up and nose-down directions, which results from the 
greater of--
    (1) A three-second movement of the longitudinal trim system at its 
normal rate for the particular flight condition with no aerodynamic load 
(or an equivalent degree of trim for airplanes that do not have a power-
operated trim system), except as limited by stops in the trim system, 
including those required by Sec. 25.655(b) for adjustable stabilizers; 
or
    (2) The maximum mistrim that can be sustained by the autopilot while 
maintaining level flight in the high speed cruising condition.
    (b) In the out-of-trim condition specified in paragraph (a) of this 
section, when the normal acceleration is varied from +1 g to the 
positive and negative values specified in paragraph (c) of this 
section--
    (1) The stick force vs. g curve must have a positive slope at any 
speed up to and including VFC/MFC; and
    (2) At speeds between VFC/MFC and 
VDF/MDF the direction of the primary longitudinal 
control force may not reverse.
    (c) Except as provided in paragraphs (d) and (e) of this section, 
compliance with the provisions of paragraph (a) of this section must be 
demonstrated in flight over the acceleration range--
    (1) -1 g to +2.5 g; or
    (2) 0 g to 2.0 g, and extrapolating by an acceptable method to -1 g 
and +2.5 g.
    (d) If the procedure set forth in paragraph (c)(2) of this section 
is used to demonstrate compliance and marginal conditions exist during 
flight test with regard to reversal of primary longitudinal control 
force, flight tests must be accomplished from the normal acceleration at 
which a marginal condition is found to exist to the applicable limit

[[Page 400]]

specified in paragraph (b)(1) of this section.
    (e) During flight tests required by paragraph (a) of this section, 
the limit maneuvering load factors prescribed in Sec. Sec. 25.333(b) 
and 25.337, and the maneuvering load factors associated with probable 
inadvertent excursions beyond the boundaries of the buffet onset 
envelopes determined under Sec. 25.251(e), need not be exceeded. In 
addition, the entry speeds for flight test demonstrations at normal 
acceleration values less than 1 g must be limited to the extent 
necessary to accomplish a recovery without exceeding VDF/
MDF.
    (f) In the out-of-trim condition specified in paragraph (a) of this 
section, it must be possible from an overspeed condition at 
VDF/MDF to produce at least 1.5 g for recovery by 
applying not more than 125 pounds of longitudinal control force using 
either the primary longitudinal control alone or the primary 
longitudinal control and the longitudinal trim system. If the 
longitudinal trim is used to assist in producing the required load 
factor, it must be shown at VDF/MDF that the 
longitudinal trim can be actuated in the airplane nose-up direction with 
the primary surface loaded to correspond to the least of the following 
airplane nose-up control forces:
    (1) The maximum control forces expected in service as specified in 
Sec. Sec. 25.301 and 25.397.
    (2) The control force required to produce 1.5 g.
    (3) The control force corresponding to buffeting or other phenomena 
of such intensity that it is a strong deterrent to further application 
of primary longitudinal control force.

[Amdt. 25-42, 43 FR 2322, Jan. 16, 1978]



                           Subpart C_Structure

                                 General



Sec. 25.301  Loads.

    (a) Strength requirements are specified in terms of limit loads (the 
maximum loads to be expected in service) and ultimate loads (limit loads 
multiplied by prescribed factors of safety). Unless otherwise provided, 
prescribed loads are limit loads.
    (b) Unless otherwise provided, the specified air, ground, and water 
loads must be placed in equilibrium with inertia forces, considering 
each item of mass in the airplane. These loads must be distributed to 
conservatively approximate or closely represent actual conditions. 
Methods used to determine load intensities and distribution must be 
validated by flight load measurement unless the methods used for 
determining those loading conditions are shown to be reliable.
    (c) If deflections under load would significantly change the 
distribution of external or internal loads, this redistribution must be 
taken into account.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970]



Sec. 25.303  Factor of safety.

    Unless otherwise specified, a factor of safety of 1.5 must be 
applied to the prescribed limit load which are considered external loads 
on the structure. When a loading condition is prescribed in terms of 
ultimate loads, a factor of safety need not be applied unless otherwise 
specified.

[Amdt. 25-23, 35 FR 5672, Apr. 8, 1970]



Sec. 25.305  Strength and deformation.

    (a) The structure must be able to support limit loads without 
detrimental permanent deformation. At any load up to limit loads, the 
deformation may not interfere with safe operation.
    (b) The structure must be able to support ultimate loads without 
failure for at least 3 seconds. However, when proof of strength is shown 
by dynamic tests simulating actual load conditions, the 3-second limit 
does not apply. Static tests conducted to ultimate load must include the 
ultimate deflections and ultimate deformation induced by the loading. 
When analytical methods are used to show compliance with the ultimate 
load strength requirements, it must be shown that--
    (1) The effects of deformation are not significant;
    (2) The deformations involved are fully accounted for in the 
analysis; or
    (3) The methods and assumptions used are sufficient to cover the 
effects of these deformations.

[[Page 401]]

    (c) Where structural flexibility is such that any rate of load 
application likely to occur in the operating conditions might produce 
transient stresses appreciably higher than those corresponding to static 
loads, the effects of this rate of application must be considered.
    (d) [Reserved]
    (e) The airplane must be designed to withstand any vibration and 
buffeting that might occur in any likely operating condition up to 
VD/MD, including stall and probable inadvertent 
excursions beyond the boundaries of the buffet onset envelope. This must 
be shown by analysis, flight tests, or other tests found necessary by 
the Administrator.
    (f) Unless shown to be extremely improbable, the airplane must be 
designed to withstand any forced structural vibration resulting from any 
failure, malfunction or adverse condition in the flight control system. 
These must be considered limit loads and must be investigated at 
airspeeds up to VC/MC.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; 
Amdt. 25-77, 57 FR 28949, June 29, 1992; Amdt. 25-86, 61 FR 5220, Feb. 
9, 1996]



Sec. 25.307  Proof of structure.

    (a) Compliance with the strength and deformation requirements of 
this subpart must be shown for each critical loading condition. 
Structural analysis may be used only if the structure conforms to that 
for which experience has shown this method to be reliable. In other 
cases, substantiating tests must be made to load levels that are 
sufficient to verify structural behavior up to loads specified in Sec. 
25.305.
    (b)-(c) [Reserved]
    (d) When static or dynamic tests are used to show compliance with 
the requirements of Sec. 25.305(b) for flight structures, appropriate 
material correction factors must be applied to the test results, unless 
the structure, or part thereof, being tested has features such that a 
number of elements contribute to the total strength of the structure and 
the failure of one element results in the redistribution of the load 
through alternate load paths.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; 
Amdt. 25-72, 55 FR 29775, July 20, 1990; 79 FR 59429, Oct. 2, 2014]

                              Flight Loads



Sec. 25.321  General.

    (a) Flight load factors represent the ratio of the aerodynamic force 
component (acting normal to the assumed longitudinal axis of the 
airplane) to the weight of the airplane. A positive load factor is one 
in which the aerodynamic force acts upward with respect to the airplane.
    (b) Considering compressibility effects at each speed, compliance 
with the flight load requirements of this subpart must be shown--
    (1) At each critical altitude within the range of altitudes selected 
by the applicant;
    (2) At each weight from the design minimum weight to the design 
maximum weight appropriate to each particular flight load condition; and
    (3) For each required altitude and weight, for any practicable 
distribution of disposable load within the operating limitations 
recorded in the Airplane Flight Manual.
    (c) Enough points on and within the boundaries of the design 
envelope must be investigated to ensure that the maximum load for each 
part of the airplane structure is obtained.
    (d) The significant forces acting on the airplane must be placed in 
equilibrium in a rational or conservative manner. The linear inertia 
forces must be considered in equilibrium with the thrust and all 
aerodynamic loads, while the angular (pitching) inertia forces must be 
considered in equilibrium with thrust and all aerodynamic moments, 
including moments due to loads on components such as tail surfaces and 
nacelles. Critical thrust values in the range from zero to maximum 
continuous thrust must be considered.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-86, 61 FR 5220, Feb. 9, 1996]

[[Page 402]]

                   Flight Maneuver and Gust Conditions



Sec. 25.331  Symmetric maneuvering conditions.

    (a) Procedure. For the analysis of the maneuvering flight conditions 
specified in paragraphs (b) and (c) of this section, the following 
provisions apply:
    (1) Where sudden displacement of a control is specified, the assumed 
rate of control surface displacement may not be less than the rate that 
could be applied by the pilot through the control system.
    (2) In determining elevator angles and chordwise load distribution 
in the maneuvering conditions of paragraphs (b) and (c) of this section, 
the effect of corresponding pitching velocities must be taken into 
account. The in-trim and out-of-trim flight conditions specified in 
Sec. 25.255 must be considered.
    (b) Maneuvering balanced conditions. Assuming the airplane to be in 
equilibrium with zero pitching acceleration, the maneuvering conditions 
A through I on the maneuvering envelope in Sec. 25.333(b) must be 
investigated.
    (c) Pitch maneuver conditions. The conditions specified in 
paragraphs (c)(1) and (2) of this section must be investigated. The 
movement of the pitch control surfaces may be adjusted to take into 
account limitations imposed by the maximum pilot effort specified by 
Sec. 25.397(b), control system stops and any indirect effect imposed by 
limitations in the output side of the control system (for example, 
stalling torque or maximum rate obtainable by a power control system.)
    (1) Maximum pitch control displacement at VA. The 
airplane is assumed to be flying in steady level flight (point 
A1, Sec. 25.333(b)) and the cockpit pitch control is 
suddenly moved to obtain extreme nose up pitching acceleration. In 
defining the tail load, the response of the airplane must be taken into 
account. Airplane loads that occur subsequent to the time when normal 
acceleration at the c.g. exceeds the positive limit maneuvering load 
factor (at point A2 in Sec. 25.333(b)), or the resulting 
tailplane normal load reaches its maximum, whichever occurs first, need 
not be considered.
    (2) Specified control displacement. A checked maneuver, based on a 
rational pitching control motion vs. time profile, must be established 
in which the design limit load factor specified in Sec. 25.337 will not 
be exceeded. Unless lesser values cannot be exceeded, the airplane 
response must result in pitching accelerations not less than the 
following:
    (i) A positive pitching acceleration (nose up) is assumed to be 
reached concurrently with the airplane load factor of 1.0 (Points 
A1 to D1, Sec. 25.333(b)). The positive 
acceleration must be equal to at least
[GRAPHIC] [TIFF OMITTED] TC28SE91.033

where--

n is the positive load factor at the speed under consideration, and V is 
          the airplane equivalent speed in knots.

    (ii) A negative pitching acceleration (nose down) is assumed to be 
reached concurrently with the positive maneuvering load factor (points 
A2 to D2, Sec. 25.333(b)). This negative pitching 
acceleration must be equal to at least
[GRAPHIC] [TIFF OMITTED] TC28SE91.034

where--

n is the positive load factor at the speed under consideration; and V is 
          the airplane equivalent speed in knots.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-46, 43 FR 50594, Oct. 30, 1978; 43 FR 
52495, Nov. 13, 1978; 43 FR 54082, Nov. 20, 1978; Amdt. 25-72, 55 FR 
29775, July 20, 1990; 55 FR 37607, Sept. 12, 1990; Amdt. 25-86, 61 FR 
5220, Feb. 9, 1996; Amdt. 25-91, 62 FR 40704, July 29, 1997]

    Effective Date Note: At 79 FR 73466, Dec. 11, 2014, Sec. 25.331 was 
amended by revising paragraph (c) introductory text and paragraph 
(c)(2), effective Feb. 9, 2015. For the convenience of the user, the 
revised text is set forth as follows:



Sec. 25.331  Symmetric maneuvering conditions.

                                * * * * *

    (c) Maneuvering pitching conditions. The following conditions must 
be investigated:

                                * * * * *

[[Page 403]]

    (2) Checked maneuver between VA and VD. Nose-up checked pitching 
maneuvers must be analyzed in which the positive limit load factor 
prescribed in Sec. 25.337 is achieved. As a separate condition, nose-
down checked pitching maneuvers must be analyzed in which a limit load 
factor of 0g is achieved. In defining the airplane loads, the flight 
deck pitch control motions described in paragraphs (c)(2)(i) through 
(iv) of this section must be used:
    (i) The airplane is assumed to be flying in steady level flight at 
any speed between VA and VD and the flight deck 
pitch control is moved in accordance with the following formula:

[delta](t) = [delta]1 sin([omega]t) for 0 <= t <=t 
max

Where--

[delta]1 = the maximum available displacement of the flight 
          deck pitch control in the initial direction, as limited by the 
          control system stops, control surface stops, or by pilot 
          effort in accordance with Sec. 25.397(b);
[delta](t) = the displacement of the flight deck pitch control as a 
          function of time. In the initial direction, [delta](t) is 
          limited to [delta]1. In the reverse direction, 
          [delta](t) may be truncated at the maximum available 
          displacement of the flight deck pitch control as limited by 
          the control system stops, control surface stops, or by pilot 
          effort in accordance with 25.397(b);
tmax = 3[pi]/2[omega];
[omega] = the circular frequency (radians/second) of the control 
          deflection taken equal to the undamped natural frequency of 
          the short period rigid mode of the airplane, with active 
          control system effects included where appropriate; but not 
          less than:
          [GRAPHIC] [TIFF OMITTED] TR11DE14.023
          
Where

V = the speed of the airplane at entry to the maneuver.
VA = the design maneuvering speed prescribed in Sec. 
          25.335(c).

    (ii) For nose-up pitching maneuvers, the complete flight deck pitch 
control displacement history may be scaled down in amplitude to the 
extent necessary to ensure that the positive limit load factor 
prescribed in Sec. 25.337 is not exceeded. For nose-down pitching 
maneuvers, the complete flight deck control displacement history may be 
scaled down in amplitude to the extent necessary to ensure that the 
normal acceleration at the center of gravity does not go below 0g.
    (iii) In addition, for cases where the airplane response to the 
specified flight deck pitch control motion does not achieve the 
prescribed limit load factors, then the following flight deck pitch 
control motion must be used:

[delta](t) = [delta]1 sin([omega]t) for 0 <= t <= 
t1
[delta](t) = [delta]1 for t1 <= t <= t2
[delta](t) = [delta]1 sin([omega][t + t1 - 
t2]) for t2 <= t <= tmax

Where--

t1 = [pi]/2[omega]
t2 = t1 + [Delta]t
tmax = t2 + [pi]/[omega];
[Delta]t = the minimum period of time necessary to allow the prescribed 
          limit load factor to be achieved in the initial direction, but 
          it need not exceed five seconds (see figure below).
          [GRAPHIC] [TIFF OMITTED] TR11DE14.024
          

[[Page 404]]


    (iv) In cases where the flight deck pitch control motion may be 
affected by inputs from systems (for example, by a stick pusher that can 
operate at high load factor as well as at 1g), then the effects of those 
systems shall be taken into account.
    (v) Airplane loads that occur beyond the following times need not be 
considered:
    (A) For the nose-up pitching maneuver, the time at which the normal 
acceleration at the center of gravity goes below 0g;
    (B) For the nose-down pitching maneuver, the time at which the 
normal acceleration at the center of gravity goes above the positive 
limit load factor prescribed in Sec. 25.337;
    (C) tmax..



Sec. 25.333  Flight maneuvering envelope.

    (a) General. The strength requirements must be met at each 
combination of airspeed and load factor on and within the boundaries of 
the representative maneuvering envelope (V-n diagram) of paragraph (b) 
of this section. This envelope must also be used in determining the 
airplane structural operating limitations as specified in Sec. 25.1501.
    (b) Maneuvering envelope.
    [GRAPHIC] [TIFF OMITTED] TC28SE91.035
    

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-86, 
61 FR 5220, Feb. 9, 1996]



Sec. 25.335  Design airspeeds.

    The selected design airspeeds are equivalent airspeeds (EAS). 
Estimated values of VS0 and VS1 must be 
conservative.
    (a) Design cruising speed, VC. For VC, the following apply:
    (1) The minimum value of VC must be sufficiently greater than VB to 
provide for inadvertent speed increases likely to occur as a result of 
severe atmospheric turbulence.
    (2) Except as provided in Sec. 25.335(d)(2), VC may not 
be less than VB + 1.32 U REF (with UREF 
as specified in Sec. 25.341(a)(5)(i)). However VC need not 
exceed the maximum speed in level flight at maximum continuous power for 
the corresponding altitude.
    (3) At altitudes where VD is limited by Mach number, VC may be 
limited to a selected Mach number.
    (b) Design dive speed, VD. VD must be selected so that VC/MC is not 
greater

[[Page 405]]

than 0.8 VD/MD, or so that the minimum speed margin between VC/MC and 
VD/MD is the greater of the following values:
    (1) From an initial condition of stabilized flight at VC/MC, the 
airplane is upset, flown for 20 seconds along a flight path 7.5[deg] 
below the initial path, and then pulled up at a load factor of 1.5g 
(0.5g acceleration increment). The speed increase occurring in this 
maneuver may be calculated if reliable or conservative aerodynamic data 
is used. Power as specified in Sec. 25.175(b)(1)(iv) is assumed until 
the pullup is initiated, at which time power reduction and the use of 
pilot controlled drag devices may be assumed;
    (2) The minimum speed margin must be enough to provide for 
atmospheric variations (such as horizontal gusts, and penetration of jet 
streams and cold fronts) and for instrument errors and airframe 
production variations. These factors may be considered on a probability 
basis. The margin at altitude where MC is limited by 
compressibility effects must not less than 0.07M unless a lower margin 
is determined using a rational analysis that includes the effects of any 
automatic systems. In any case, the margin may not be reduced to less 
than 0.05M.
    (c) Design maneuvering speed VA. For VA, the following apply:
    (1) VA may not be less than VS1 [radic]n where--
    (i) n is the limit positive maneuvering load factor at VC; and
    (ii) VS1 is the stalling speed with flaps retracted.
    (2) VA and VS must be evaluated at the design weight and altitude 
under consideration.
    (3) VA need not be more than VC or the speed at which the positive 
CN max curve intersects the positive maneuver load factor line, 
whichever is less.
    (d) Design speed for maximum gust intensity, VB.
    (1) VB may not be less than
    [GRAPHIC] [TIFF OMITTED] TR09FE96.016
    
where--

VS1=the 1-g stalling speed based on CNAmax with 
          the flaps retracted at the particular weight under 
          consideration;
Vc=design cruise speed (knots equivalent airspeed);
Uref=the reference gust velocity (feet per second equivalent 
          airspeed) from Sec. 25.341(a)(5)(i);
w=average wing loading (pounds per square foot) at the particular weight 
          under consideration.
          [GRAPHIC] [TIFF OMITTED] TR09FE96.017
          
[rho]=density of air (slugs/ft\3\);
c=mean geometric chord of the wing (feet);
g=acceleration due to gravity (ft/sec\2\);
a=slope of the airplane normal force coefficient curve, CNA 
          per radian;

    (2) At altitudes where VC is limited by Mach number--
    (i) VB may be chosen to provide an optimum margin between 
low and high speed buffet boundaries; and,
    (ii) VB need not be greater than VC.
    (e) Design flap speeds, VF. For VF, the following apply:
    (1) The design flap speed for each flap position (established in 
accordance with Sec. 25.697(a)) must be sufficiently greater than the 
operating speed recommended for the corresponding stage of flight 
(including balked landings) to allow for probable variations in control 
of airspeed and for transition from one flap position to another.
    (2) If an automatic flap positioning or load limiting device is 
used, the speeds and corresponding flap positions programmed or allowed 
by the device may be used.
    (3) VF may not be less than--
    (i) 1.6 VS1 with the flaps in takeoff position at maximum 
takeoff weight;
    (ii) 1.8 VS1 with the flaps in approach position at 
maximum landing weight, and
    (iii) 1.8 VS0 with the flaps in landing position at 
maximum landing weight.
    (f) Design drag device speeds, VDD. The selected design speed for 
each drag device must be sufficiently greater than the speed recommended 
for the operation of the device to allow for probable variations in 
speed control. For drag devices intended for use in high speed descents, 
VDD may not be less

[[Page 406]]

than VD. When an automatic drag device positioning or load limiting 
means is used, the speeds and corresponding drag device positions 
programmed or allowed by the automatic means must be used for design.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-86, 61 FR 5220, Feb. 9, 1996; Amdt. 
25-91, 62 FR 40704, July 29, 1997]



Sec. 25.337  Limit maneuvering load factors.

    (a) Except where limited by maximum (static) lift coefficients, the 
airplane is assumed to be subjected to symmetrical maneuvers resulting 
in the limit maneuvering load factors prescribed in this section. 
Pitching velocities appropriate to the corresponding pull-up and steady 
turn maneuvers must be taken into account.
    (b) The positive limit maneuvering load factor n for any speed up to 
Vn may not be less than 2.1+24,000/ (W +10,000) except that n may not be 
less than 2.5 and need not be greater than 3.8--where W is the design 
maximum takeoff weight.
    (c) The negative limit maneuvering load factor--
    (1) May not be less than -1.0 at speeds up to VC; and
    (2) Must vary linearly with speed from the value at VC to zero at 
VD.
    (d) Maneuvering load factors lower than those specified in this 
section may be used if the airplane has design features that make it 
impossible to exceed these values in flight.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970]



Sec. 25.341  Gust and turbulence loads.

    (a) Discrete Gust Design Criteria. The airplane is assumed to be 
subjected to symmetrical vertical and lateral gusts in level flight. 
Limit gust loads must be determined in accordance with the provisions:
    (1) Loads on each part of the structure must be determined by 
dynamic analysis. The analysis must take into account unsteady 
aerodynamic characteristics and all significant structural degrees of 
freedom including rigid body motions.
    (2) The shape of the gust must be:
    [GRAPHIC] [TIFF OMITTED] TR09FE96.018
    
for 0 <=s <=2H
where--
s=distance penetrated into the gust (feet);
Uds=the design gust velocity in equivalent airspeed specified 
          in paragraph (a)(4) of this section; and
H=the gust gradient which is the distance (feet) parallel to the 
          airplane's flight path for the gust to reach its peak 
          velocity.

    (3) A sufficient number of gust gradient distances in the range 30 
feet to 350 feet must be investigated to find the critical response for 
each load quantity.
    (4) The design gust velocity must be:
    [GRAPHIC] [TIFF OMITTED] TR09FE96.019
    
where--

Uref=the reference gust velocity in equivalent airspeed 
          defined in paragraph (a)(5) of this section.
Fg=the flight profile alleviation factor defined in paragraph 
          (a)(6) of this section.

    (5) The following reference gust velocities apply:
    (i) At the airplane design speed VC: Positive and 
negative gusts with reference gust velocities of 56.0 ft/sec EAS must be 
considered at sea level. The reference gust velocity may be reduced 
linearly from 56.0 ft/sec EAS at sea level to 44.0 ft/sec EAS at 15000 
feet. The reference gust velocity may be further reduced linearly from 
44.0 ft/sec EAS at 15000 feet to 26.0 ft/sec EAS at 50000 feet.
    (ii) At the airplane design speed VD: The reference gust 
velocity must be 0.5 times the value obtained under Sec. 
25.341(a)(5)(i).
    (6) The flight profile alleviation factor, Fg, must be 
increased linearly from the sea level value to a value of 1.0 at the 
maximum operating altitude defined in Sec. 25.1527. At sea level, the 
flight profile alleviation factor is determined by the following 
equation:

[[Page 407]]

[GRAPHIC] [TIFF OMITTED] TN08MR96.004

Zmo=Maximum operating altitude defined in Sec. 25.1527.

    (7) When a stability augmentation system is included in the 
analysis, the effect of any significant system nonlinearities should be 
accounted for when deriving limit loads from limit gust conditions.
    (b) Continuous Gust Design Criteria. The dynamic response of the 
airplane to vertical and lateral continuous turbulence must be taken 
into account. The continuous gust design criteria of appendix G of this 
part must be used to establish the dynamic response unless more rational 
criteria are shown.

[Doc. No. 27902, 61 FR 5221, Feb. 9, 1996; 61 FR 9533, Mar. 8, 1996]

    Effective Date Note: At 79 FR 73467, Dec. 11, 2014, Sec. 25.341 was 
amended by revising paragraphs (a)(5)(i), (a)(6), and (b), and by adding 
paragraph (c), effective Feb. 9, 2015. For the convenience of the user, 
the revised and added text is set forth as follows:



Sec. 25.341  Gust and turbulence loads.

    (a) * * *

                                * * * * *

    (5) * * *
    (i) At airplane speeds between VB and VC: 
Positive and negative gusts with reference gust velocities of 56.0 ft/
sec EAS must be considered at sea level. The reference gust velocity may 
be reduced linearly from 56.0 ft/sec EAS at sea level to 44.0 ft/sec EAS 
at 15,000 feet. The reference gust velocity may be further reduced 
linearly from 44.0 ft/sec EAS at 15,000 feet to 20.86 ft/sec EAS at 
60,000 feet.

                                * * * * *

    (6) * * *

Zmo = Maximum operating altitude defined in Sec. 25.1527 
          (feet).

                                * * * * *

    (b) Continuous turbulence design criteria. The dynamic response of 
the airplane to vertical and lateral continuous turbulence must be taken 
into account. The dynamic analysis must take into account unsteady 
aerodynamic characteristics and all significant structural degrees of 
freedom including rigid body motions. The limit loads must be determined 
for all critical altitudes, weights, and weight distributions as 
specified in Sec. 25.321(b), and all critical speeds within the ranges 
indicated in Sec. 25.341(b)(3).
    (1) Except as provided in paragraphs (b)(4) and (5) of this section, 
the following equation must be used:

PL = PL-1g  U[sigma]A

Where--

PL = limit load;
PL-1g = steady 1g load for the condition;
A = ratio of root-mean-square incremental load for the condition to 
          root-mean-square turbulence velocity; and
U[sigma] = limit turbulence intensity in true airspeed, specified in 
          paragraph (b)(3) of this section.

    (2) Values of A must be determined according to the following 
formula:
[GRAPHIC] [TIFF OMITTED] TR11DE14.025


[[Page 408]]


Where--

H([Omega]) = the frequency response function, determined by dynamic 
          analysis, that relates the loads in the aircraft structure to 
          the atmospheric turbulence; and
[Phi]([Omega]) = normalized power spectral density of atmospheric 
          turbulence given by--
          [GRAPHIC] [TIFF OMITTED] TR11DE14.026
          
Where--

[Omega] = reduced frequency, radians per foot; and
L = scale of turbulence = 2,500 ft.

    (3) The limit turbulence intensities, U[sigma], in feet per second 
true airspeed required for compliance with this paragraph are--
    (i) At airplane speeds between VB and VC: 
U[sigma] = U[sigma]ref Fg

Where--

U[sigma]ref is the reference turbulence intensity that varies linearly 
          with altitude from 90 fps (TAS) at sea level to 79 fps (TAS) 
          at 24,000 feet and is then constant at 79 fps (TAS) up to the 
          altitude of 60,000 feet.
Fg is the flight profile alleviation factor defined in 
          paragraph (a)(6) of this section;

    (ii) At speed VD: U[sigma] is equal to \1/2\ the values 
obtained under paragraph (b)(3)(i) of this section.
    (iii) At speeds between VC and VD: U[sigma] is 
equal to a value obtained by linear interpolation.
    (iv) At all speeds, both positive and negative incremental loads due 
to continuous turbulence must be considered.
    (4) When an automatic system affecting the dynamic response of the 
airplane is included in the analysis, the effects of system non-
linearities on loads at the limit load level must be taken into account 
in a realistic or conservative manner.
    (5) If necessary for the assessment of loads on airplanes with 
significant non-linearities, it must be assumed that the turbulence 
field has a root-mean-square velocity equal to 40 percent of the 
U[sigma] values specified in paragraph (b)(3) of this section. The value 
of limit load is that load with the same probability of exceedance in 
the turbulence field as AU[sigma] of the same load quantity in a linear 
approximated model.
    (c) Supplementary gust conditions for wing-mounted engines. For 
airplanes equipped with wing-mounted engines, the engine mounts, pylons, 
and wing supporting structure must be designed for the maximum response 
at the nacelle center of gravity derived from the following dynamic gust 
conditions applied to the airplane:
    (1) A discrete gust determined in accordance with Sec. 25.341(a) at 
each angle normal to the flight path, and separately,
    (2) A pair of discrete gusts, one vertical and one lateral. The 
length of each of these gusts must be independently tuned to the maximum 
response in accordance with Sec. 25.341(a). The penetration of the 
airplane in the combined gust field and the phasing of the vertical and 
lateral component gusts must be established to develop the maximum 
response to the gust pair. In the absence of a more rational analysis, 
the following formula must be used for each of the maximum engine loads 
in all six degrees of freedom:
[GRAPHIC] [TIFF OMITTED] TR11DE14.027

Where--

PL = limit load;
PL-1g = steady 1g load for the condition;
LV = peak incremental response load due to a vertical gust 
          according to Sec. 25.341(a); and
LL = peak incremental response load due to a lateral gust 
          according to Sec. 25.341(a).



Sec. 25.343  Design fuel and oil loads.

    (a) The disposable load combinations must include each fuel and oil 
load in the range from zero fuel and oil to the selected maximum fuel 
and oil load. A structural reserve fuel condition, not exceeding 45 
minutes of fuel under the

[[Page 409]]

operating conditions in Sec. 25.1001(e) and (f), as applicable, may be 
selected.
    (b) If a structural reserve fuel condition is selected, it must be 
used as the minimum fuel weight condition for showing compliance with 
the flight load requirements as prescribed in this subpart. In 
addition--
    (1) The structure must be designed for a condition of zero fuel and 
oil in the wing at limit loads corresponding to--
    (i) A maneuvering load factor of +2.25; and
    (ii) The gust conditions of Sec. 25.341(a) but assuming 85% of the 
design velocities prescribed in Sec. 25.341(a)(4).
    (2) Fatigue evaluation of the structure must account for any 
increase in operating stresses resulting from the design condition of 
paragraph (b)(1) of this section; and
    (3) The flutter, deformation, and vibration requirements must also 
be met with zero fuel.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-18, 
33 FR 12226, Aug. 30, 1968; Amdt. 25-72, 55 FR 37607, Sept. 12, 1990; 
Amdt. 25-86, 61 FR 5221, Feb. 9, 1996]

    Effective Date Note: At 79 FR 73468, Dec. 11, 2014, Sec. 25.343 was 
amended by revising paragraph (b)(1)(ii), effective Feb. 9, 2015. For 
the convenience of the user, the revised text is set forth as follows:



Sec. 25.343  Design fuel and oil loads.

                                * * * * *

    (b) * * *
    (1) * * *
    (ii) The gust and turbulence conditions of Sec. 25.341(a) and (b), 
but assuming 85% of the gust velocities prescribed in Sec. 25.341(a)(4) 
and 85% of the turbulence intensities prescribed in Sec. 25.341(b)(3).

                                * * * * *



Sec. 25.345  High lift devices.

    (a) If wing flaps are to be used during takeoff, approach, or 
landing, at the design flap speeds established for these stages of 
flight under Sec. 25.335(e) and with the wing flaps in the 
corresponding positions, the airplane is assumed to be subjected to 
symmetrical maneuvers and gusts. The resulting limit loads must 
correspond to the conditions determined as follows:
    (1) Maneuvering to a positive limit load factor of 2.0; and
    (2) Positive and negative gusts of 25 ft/sec EAS acting normal to 
the flight path in level flight. Gust loads resulting on each part of 
the structure must be determined by rational analysis. The analysis must 
take into account the unsteady aerodynamic characteristics and rigid 
body motions of the aircraft. The shape of the gust must be as described 
in Sec. 25.341(a)(2) except that--

Uds=25 ft/sec EAS;
H=12.5 c; and
c=mean geometric chord of the wing (feet).

    (b) The airplane must be designed for the conditions prescribed in 
paragraph (a) of this section, except that the airplane load factor need 
not exceed 1.0, taking into account, as separate conditions, the effects 
of--
    (1) Propeller slipstream corresponding to maximum continuous power 
at the design flap speeds VF, and with takeoff power at not less than 
1.4 times the stalling speed for the particular flap position and 
associated maximum weight; and
    (2) A head-on gust of 25 feet per second velocity (EAS).
    (c) If flaps or other high lift devices are to be used in en route 
conditions, and with flaps in the appropriate position at speeds up to 
the flap design speed chosen for these conditions, the airplane is 
assumed to be subjected to symmetrical maneuvers and gusts within the 
range determined by--
    (1) Maneuvering to a positive limit load factor as prescribed in 
Sec. 25.337(b); and
    (2) The discrete vertical gust criteria in Sec. 25.341(a).
    (d) The airplane must be designed for a maneuvering load factor of 
1.5 g at the maximum take-off weight with the wing-flaps and similar 
high lift devices in the landing configurations.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 
43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55 FR 37607, Sept. 17, 1990; 
Amdt. 25-86, 61 FR 5221, Feb. 9, 1996; Amdt. 25-91, 62 FR 40704, July 
29, 1997]

    Effective Date Note: At 79 FR 73468, Dec. 11, 2014, Sec. 25.345 was 
amended by revising paragraph (c)(2), effective Feb. 9, 2015. For the 
convenience of the user, the revised text is set forth as follows:

[[Page 410]]



Sec. 25.345  High lift devices.

                                * * * * *

    (c) * * *
    (2) The vertical gust and turbulence conditions prescribed in Sec. 
25.341(a) and (b).

                                * * * * *



Sec. 25.349  Rolling conditions.

    The airplane must be designed for loads resulting from the rolling 
conditions specified in paragraphs (a) and (b) of this section. 
Unbalanced aerodynamic moments about the center of gravity must be 
reacted in a rational or conservative manner, considering the principal 
masses furnishing the reacting inertia forces.
    (a) Maneuvering. The following conditions, speeds, and aileron 
deflections (except as the deflections may be limited by pilot effort) 
must be considered in combination with an airplane load factor of zero 
and of two-thirds of the positive maneuvering factor used in design. In 
determining the required aileron deflections, the torsional flexibility 
of the wing must be considered in accordance with Sec. 25.301(b):
    (1) Conditions corresponding to steady rolling velocities must be 
investigated. In addition, conditions corresponding to maximum angular 
acceleration must be investigated for airplanes with engines or other 
weight concentrations outboard of the fuselage. For the angular 
acceleration conditions, zero rolling velocity may be assumed in the 
absence of a rational time history investigation of the maneuver.
    (2) At VA, a sudden deflection of the aileron to the stop is 
assumed.
    (3) At VC, the aileron deflection must be that required to produce a 
rate of roll not less than that obtained in paragraph (a)(2) of this 
section.
    (4) At VD, the aileron deflection must be that required to produce a 
rate of roll not less than one-third of that in paragraph (a)(2) of this 
section.
    (b) Unsymmetrical gusts. The airplane is assumed to be subjected to 
unsymmetrical vertical gusts in level flight. The resulting limit loads 
must be determined from either the wing maximum airload derived directly 
from Sec. 25.341(a), or the wing maximum airload derived indirectly 
from the vertical load factor calculated from Sec. 25.341(a). It must 
be assumed that 100 percent of the wing air load acts on one side of the 
airplane and 80 percent of the wing air load acts on the other side.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-86, 61 FR 5222, Feb. 9, 1996; Amdt. 
25-94, 63 FR 8848, Feb. 23, 1998]



Sec. 25.351  Yaw maneuver conditions.

    The airplane must be designed for loads resulting from the yaw 
maneuver conditions specified in paragraphs (a) through (d) of this 
section at speeds from VMC to VD. Unbalanced 
aerodynamic moments about the center of gravity must be reacted in a 
rational or conservative manner considering the airplane inertia forces. 
In computing the tail loads the yawing velocity may be assumed to be 
zero.
    (a) With the airplane in unaccelerated flight at zero yaw, it is 
assumed that the cockpit rudder control is suddenly displaced to achieve 
the resulting rudder deflection, as limited by:
    (1) The control system on control surface stops; or
    (2) A limit pilot force of 300 pounds from VMC to 
VA and 200 pounds from VC/MC to 
VD/MD, with a linear variation between 
VA and VC/MC.
    (b) With the cockpit rudder control deflected so as always to 
maintain the maximum rudder deflection available within the limitations 
specified in paragraph (a) of this section, it is assumed that the 
airplane yaws to the overswing sideslip angle.
    (c) With the airplane yawed to the static equilibrium sideslip 
angle, it is assumed that the cockpit rudder control is held so as to 
achieve the maximum rudder deflection available within the limitations 
specified in paragraph (a) of this section.
    (d) With the airplane yawed to the static equilibrium sideslip angle 
of paragraph (c) of this section, it is assumed that the cockpit rudder 
control is suddenly returned to neutral.

[Amdt. 25-91, 62 FR 40704, July 29, 1997]

[[Page 411]]

                        Supplementary Conditions



Sec. 25.361  Engine torque.

    (a) Each engine mount and its supporting structure must be designed 
for the effects of--
    (1) A limit engine torque corresponding to takeoff power and 
propeller speed acting simultaneously with 75 percent of the limit loads 
from flight condition A of Sec. 25.333(b);
    (2) A limit torque corresponding to the maximum continuous power and 
propeller speed, acting simultaneously with the limit loads from flight 
condition A of Sec. 25.333(b); and
    (3) For turbopropeller installations, in addition to the conditions 
specified in paragraphs (a)(1) and (2) of this section, a limit engine 
torque corresponding to takeoff power and propeller speed, multiplied by 
a factor accounting for propeller control system malfunction, including 
quick feathering, acting simultaneously with 1g level flight loads. In 
the absence of a rational analysis, a factor of 1.6 must be used.
    (b) For turbine engine installations, the engine mounts and 
supporting structure must be designed to withstand each of the 
following:
    (1) A limit engine torque load imposed by sudden engine stoppage due 
to malfunction or structural failure (such as compressor jamming).
    (2) A limit engine torque load imposed by the maximum acceleration 
of the engine.
    (c) The limit engine torque to be considered under paragraph (a) of 
this section must be obtained by multiplying mean torque for the 
specified power and speed by a factor of--
    (1) 1.25 for turbopropeller installations;
    (2) 1.33 for reciprocating engines with five or more cylinders; or
    (3) Two, three, or four, for engines with four, three, or two 
cylinders, respectively.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 
25-72, 55 FR 29776, July 20, 1990]

    Effective Date Note: At 79 FR 73468, Dec. 11, 2014, Sec. 25.361 was 
revised, effective Feb. 9, 2015. For the convenience of the user, the 
revised text is set forth as follows:



Sec. 25.361  Engine and auxiliary power unit torque.

    (a) For engine installations--
    (1) Each engine mount, pylon, and adjacent supporting airframe 
structures must be designed for the effects of--
    (i) A limit engine torque corresponding to takeoff power/thrust and, 
if applicable, corresponding propeller speed, acting simultaneously with 
75% of the limit loads from flight condition A of Sec. 25.333(b);
    (ii) A limit engine torque corresponding to the maximum continuous 
power/thrust and, if applicable, corresponding propeller speed, acting 
simultaneously with the limit loads from flight condition A of Sec. 
25.333(b); and
    (iii) For turbopropeller installations only, in addition to the 
conditions specified in paragraphs (a)(1)(i) and (ii) of this section, a 
limit engine torque corresponding to takeoff power and propeller speed, 
multiplied by a factor accounting for propeller control system 
malfunction, including quick feathering, acting simultaneously with 1g 
level flight loads. In the absence of a rational analysis, a factor of 
1.6 must be used.
    (2) The limit engine torque to be considered under paragraph (a)(1) 
of this section must be obtained by--
    (i) For turbopropeller installations, multiplying mean engine torque 
for the specified power/thrust and speed by a factor of 1.25;
    (ii) For other turbine engines, the limit engine torque must be 
equal to the maximum accelerating torque for the case considered.
    (3) The engine mounts, pylons, and adjacent supporting airframe 
structure must be designed to withstand 1g level flight loads acting 
simultaneously with the limit engine torque loads imposed by each of the 
following conditions to be considered separately:
    (i) Sudden maximum engine deceleration due to malfunction or 
abnormal condition; and
    (ii) The maximum acceleration of engine.
    (b) For auxiliary power unit installations, the power unit mounts 
and adjacent supporting airframe structure must be designed to withstand 
1g level flight loads acting simultaneously with the limit torque loads 
imposed by each of the following conditions to be considered separately:
    (1) Sudden maximum auxiliary power unit deceleration due to 
malfunction, abnormal condition, or structural failure; and
    (2) The maximum acceleration of the auxiliary power unit.



Sec. 25.362  Engine failure loads.

    (a) For engine mounts, pylons, and adjacent supporting airframe 
structure, an ultimate loading condition must be considered that 
combines 1g flight loads with the most critical

[[Page 412]]

transient dynamic loads and vibrations, as determined by dynamic 
analysis, resulting from failure of a blade, shaft, bearing or bearing 
support, or bird strike event. Any permanent deformation from these 
ultimate load conditions must not prevent continued safe flight and 
landing.
    (b) The ultimate loads developed from the conditions specified in 
paragraph (a) of this section are to be--
    (1) Multiplied by a factor of 1.0 when applied to engine mounts and 
pylons; and
    (2) Multiplied by a factor of 1.25 when applied to adjacent 
supporting airframe structure.

[79 FR 73468, Dec. 11, 2014]

    Effective Date Note: At 79 FR 73468, Dec. 11, 2014, Sec. 25.362 was 
added, effective Feb. 9, 2015.



Sec. 25.363  Side load on engine and auxiliary power unit mounts.

    (a) Each engine and auxiliary power unit mount and its supporting 
structure must be designed for a limit load factor in lateral direction, 
for the side load on the engine and auxiliary power unit mount, at least 
equal to the maximum load factor obtained in the yawing conditions but 
not less than--
    (1) 1.33; or
    (2) One-third of the limit load factor for flight condition A as 
prescribed in Sec. 25.333(b).
    (b) The side load prescribed in paragraph (a) of this section may be 
assumed to be independent of other flight conditions.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-91, 62 FR 40704, July 29, 1997]



Sec. 25.365  Pressurized compartment loads.

    For airplanes with one or more pressurized compartments the 
following apply:
    (a) The airplane structure must be strong enough to withstand the 
flight loads combined with pressure differential loads from zero up to 
the maximum relief valve setting.
    (b) The external pressure distribution in flight, and stress 
concentrations and fatigue effects must be accounted for.
    (c) If landings may be made with the compartment pressurized, 
landing loads must be combined with pressure differential loads from 
zero up to the maximum allowed during landing.
    (d) The airplane structure must be designed to be able to withstand 
the pressure differential loads corresponding to the maximum relief 
valve setting multiplied by a factor of 1.33 for airplanes to be 
approved for operation to 45,000 feet or by a factor of 1.67 for 
airplanes to be approved for operation above 45,000 feet, omitting other 
loads.
    (e) Any structure, component or part, inside or outside a 
pressurized compartment, the failure of which could interfere with 
continued safe flight and landing, must be designed to withstand the 
effects of a sudden release of pressure through an opening in any 
compartment at any operating altitude resulting from each of the 
following conditions:
    (1) The penetration of the compartment by a portion of an engine 
following an engine disintegration;
    (2) Any opening in any pressurized compartment up to the size 
Ho in square feet; however, small compartments may be 
combined with an adjacent pressurized compartment and both considered as 
a single compartment for openings that cannot reasonably be expected to 
be confined to the small compartment. The size Ho must be 
computed by the following formula:

Ho=PAs

where,

Ho=Maximum opening in square feet, need not exceed 20 square 
          feet.
P=(As/6240)+.024
As=Maximum cross-sectional area of the pressurized shell 
          normal to the longitudinal axis, in square feet; and

    (3) The maximum opening caused by airplane or equipment failures not 
shown to be extremely improbable.
    (f) In complying with paragraph (e) of this section, the fail-safe 
features of the design may be considered in determining the probability 
of failure or penetration and probable size of openings, provided that 
possible improper operation of closure devices and inadvertent door 
openings are also considered. Furthermore, the resulting differential 
pressure loads must be combined in a rational and conservative

[[Page 413]]

manner with 1-g level flight loads and any loads arising from emergency 
depressurization conditions. These loads may be considered as ultimate 
conditions; however, any deformations associated with these conditions 
must not interfere with continued safe flight and landing. The pressure 
relief provided by intercompartment venting may also be considered.
    (g) Bulkheads, floors, and partitions in pressurized compartments 
for occupants must be designed to withstand the conditions specified in 
paragraph (e) of this section. In addition, reasonable design 
precautions must be taken to minimize the probability of parts becoming 
detached and injuring occupants while in their seats.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-54, 
45 FR 60172, Sept. 11, 1980; Amdt. 25-71, 55 FR 13477, Apr. 10, 1990; 
Amdt. 25-72, 55 FR 29776, July 20, 1990; Amdt. 25-87, 61 FR 28695, June 
5, 1996]



Sec. 25.367  Unsymmetrical loads due to engine failure.

    (a) The airplane must be designed for the unsymmetrical loads 
resulting from the failure of the critical engine. Turbopropeller 
airplanes must be designed for the following conditions in combination 
with a single malfunction of the propeller drag limiting system, 
considering the probable pilot corrective action on the flight controls:
    (1) At speeds between VMC and VD, the loads resulting from power 
failure because of fuel flow interruption are considered to be limit 
loads.
    (2) At speeds between VMC and VC, the loads resulting from the 
disconnection of the engine compressor from the turbine or from loss of 
the turbine blades are considered to be ultimate loads.
    (3) The time history of the thrust decay and drag build-up occurring 
as a result of the prescribed engine failures must be substantiated by 
test or other data applicable to the particular engine-propeller 
combination.
    (4) The timing and magnitude of the probable pilot corrective action 
must be conservatively estimated, considering the characteristics of the 
particular engine-propeller-airplane combination.
    (b) Pilot corrective action may be assumed to be initiated at the 
time maximum yawing velocity is reached, but not earlier than two 
seconds after the engine failure. The magnitude of the corrective action 
may be based on the control forces specified in Sec. 25.397(b) except 
that lower forces may be assumed where it is shown by anaylsis or test 
that these forces can control the yaw and roll resulting from the 
prescribed engine failure conditions.



Sec. 25.371  Gyroscopic loads.

    The structure supporting any engine or auxiliary power unit must be 
designed for the loads including the gyroscopic loads arising from the 
conditions specified in Sec. Sec. 25.331, 25.341(a), 25.349, 25.351, 
25.473, 25.479, and 25.481, with the engine or auxiliary power unit at 
the maximum rpm appropriate to the condition. For the purposes of 
compliance with this section, the pitch maneuver in Sec. 25.331(c)(1) 
must be carried out until the positive limit maneuvering load factor 
(point A2 in Sec. 25.333(b)) is reached.

[Amdt. 25-91, 62 FR 40704, July 29, 1997]

    Effective Date Note: At 79 FR 73468, Dec. 11, 2014, Sec. 25.371 was 
revised, effective Feb. 9, 2015. For the convenience of the user, the 
revised text is set forth as follows:



Sec. 25.371  Gyroscopic loads.

    The structure supporting any engine or auxiliary power unit must be 
designed for the loads, including gyroscopic loads, arising from the 
conditions specified in Sec. Sec. 25.331, 25.341, 25.349, 25.351, 
25.473, 25.479, and 25.481, with the engine or auxiliary power unit at 
the maximum rotating speed appropriate to the condition. For the 
purposes of compliance with this paragraph, the pitch maneuver in Sec. 
25.331(c)(1) must be carried out until the positive limit maneuvering 
load factor (point A2 in Sec. 25.333(b)) is reached.



Sec. 25.373  Speed control devices.

    If speed control devices (such as spoilers and drag flaps) are 
installed for use in en route conditions--
    (a) The airplane must be designed for the symmetrical maneuvers 
prescribed in Sec. 25.333 and Sec. 25.337, the yawing maneuvers 
prescribed in Sec. 25.351, and the vertical and later gust conditions 
prescribed in Sec. 25.341(a), at each setting and the maximum speed 
associated with that setting; and

[[Page 414]]

    (b) If the device has automatic operating or load limiting features, 
the airplane must be designed for the maneuver and gust conditions 
prescribed in paragraph (a) of this section, at the speeds and 
corresponding device positions that the mechanism allows.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29776, July 20, 1990; Amdt. 25-86, 61 FR 5222, Feb. 9, 1996]

    Effective Date Note: At 79 FR 73468, Dec. 11, 2014, Sec. 25.373 was 
amended by revising paragraph (a), effective Feb. 9, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.373  Speed control devices.

                                * * * * *

    (a) The airplane must be designed for the symmetrical maneuvers 
prescribed in Sec. Sec. 25.333 and 25.337, the yawing maneuvers in 
Sec. 25.351, and the vertical and lateral gust and turbulence 
conditions prescribed in Sec. 25.341(a) and (b) at each setting and the 
maximum speed associated with that setting; and

                                * * * * *

                    Control Surface and System Loads



Sec. 25.391  Control surface loads: General.

    The control surfaces must be designed for the limit loads resulting 
from the flight conditions in Sec. Sec. 25.331, 25.341(a), 25.349 and 
25.351 and the ground gust conditions in Sec. 25.415, considering the 
requirements for--
    (a) Loads parallel to hinge line, in Sec. 25.393;
    (b) Pilot effort effects, in Sec. 25.397;
    (c) Trim tab effects, in Sec. 25.407;
    (d) Unsymmetrical loads, in Sec. 25.427; and
    (e) Auxiliary aerodynamic surfaces, in Sec. 25.445.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-86, 
61 FR 5222, Feb. 9, 1996]

    Effective Date Note: At 79 FR 73468, Dec. 11, 2014, Sec. 25.391 was 
amended by revising the introductory text, effective Feb. 9, 2015. For 
the convenience of the user, the revised text is set forth as follows:



Sec. 25.391  Control surface loads: General.

    The control surfaces must be designed for the limit loads resulting 
from the flight conditions in Sec. Sec. 25.331, 25.341(a) and (b), 
25.349, and 25.351, considering the requirements for--

                                * * * * *



Sec. 25.393  Loads parallel to hinge line.

    (a) Control surfaces and supporting hinge brackets must be designed 
for inertia loads acting parallel to the hinge line.
    (b) In the absence of more rational data, the inertia loads may be 
assumed to be equal to KW, where--
    (1) K=24 for vertical surfaces;
    (2) K=12 for horizontal surfaces; and
    (3) W=weight of the movable surfaces.



Sec. 25.395  Control system.

    (a) Longitudinal, lateral, directional, and drag control system and 
their supporting structures must be designed for loads corresponding to 
125 percent of the computed hinge moments of the movable control surface 
in the conditions prescribed in Sec. 25.391.
    (b) The system limit loads, except the loads resulting from ground 
gusts, need not exceed the loads that can be produced by the pilot (or 
pilots) and by automatic or power devices operating the controls.
    (c) The loads must not be less than those resulting from application 
of the minimum forces prescribed in Sec. 25.397(c).

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5672, Apr. 8, 1970; Amdt. 25-72, 55 FR 29776, July 20, 1990]

    Effective Date Note: At 79 FR 73468, Dec. 11, 2014, Sec. 25.395 was 
amended by revising paragraph (b), effective Feb. 9, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.395  Control system.

                                * * * * *

    (b) The system limit loads of paragraph (a) of this section need not 
exceed the loads that can be produced by the pilot (or pilots) and by 
automatic or power devices operating the controls.

                                * * * * *



Sec. 25.397  Control system loads.

    (a) General. The maximum and minimum pilot forces, specified in 
paragraph (c) of this section, are assumed

[[Page 415]]

to act at the appropriate control grips or pads (in a manner simulating 
flight conditions) and to be reacted at the attachment of the control 
system to the control surface horn.
    (b) Pilot effort effects. In the control surface flight loading 
condition, the air loads on movable surfaces and the corresponding 
deflections need not exceed those that would result in flight from the 
application of any pilot force within the ranges specified in paragraph 
(c) of this section. Two-thirds of the maximum values specified for the 
aileron and elevator may be used if control surface hinge moments are 
based on reliable data. In applying this criterion, the effects of servo 
mechanisms, tabs, and automatic pilot systems, must be considered.
    (c) Limit pilot forces and torques. The limit pilot forces and 
torques are as follows:

------------------------------------------------------------------------
                                    Maximum forces or  Minimum forces or
             Control                     torques            torques
------------------------------------------------------------------------
Aileron:
  Stick..........................  100 lbs...........  40 lbs.
  Wheel \1\......................  80 D in.-lbs \2\..  40 D in.-lbs.
Elevator:
  Stick..........................  250 lbs...........  100 lbs.
  Wheel (symmetrical)............  300 lbs...........  100 lbs.
  Wheel (unsymmetrical) \3\......  ..................  100 lbs.
Rudder...........................  300 lbs...........  130 lbs.
------------------------------------------------------------------------
\1\ The critical parts of the aileron control system must be designed
  for a single tangential force with a limit value equal to 1.25 times
  the couple force determined from these criteria.
\2\ D=wheel diameter (inches).
\3\ The unsymmetrical forces must be applied at one of the normal
  handgrip points on the periphery of the control wheel.


[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55466, Dec. 20, 1976; Amdt. 25-72, 55 FR 29776, July 20, 1990]



Sec. 25.399  Dual control system.

    (a) Each dual control system must be designed for the pilots 
operating in opposition, using individual pilot forces not less than--
    (1) 0.75 times those obtained under Sec. 25.395; or
    (2) The minimum forces specified in Sec. 25.397(c).
    (b) The control system must be designed for pilot forces applied in 
the same direction, using individual pilot forces not less than 0.75 
times those obtained under Sec. 25.395.



Sec. 25.405  Secondary control system.

    Secondary controls, such as wheel brake, spoiler, and tab controls, 
must be designed for the maximum forces that a pilot is likely to apply 
to those controls. The following values may be used:

             Pilot Control Force Limits (Secondary Controls)
------------------------------------------------------------------------
                  Control                        Limit pilot forces
------------------------------------------------------------------------
Miscellaneous:
  *Crank, wheel, or lever.................  ((1 + R) / 3) x 50 lbs., but
                                             not less than 50 lbs. nor
                                             more than 150 lbs.
                                             (R=radius). (Applicable to
                                             any angle within 20[deg] of
                                             plane of control).
  Twist...................................  133 in.-lbs.
  Push-pull...............................  To be chosen by applicant.
------------------------------------------------------------------------
*Limited to flap, tab, stabilizer, spoiler, and landing gear operation
  controls.



Sec. 25.407  Trim tab effects.

    The effects of trim tabs on the control surface design conditions 
must be accounted for only where the surface loads are limited by 
maximum pilot effort. In these cases, the tabs are considered to be 
deflected in the direction that would assist the pilot, and the 
deflections are--
    (a) For elevator trim tabs, those required to trim the airplane at 
any point within the positive portion of the pertinent flight envelope 
in Sec. 25.333(b), except as limited by the stops; and
    (b) For aileron and rudder trim tabs, those required to trim the 
airplane in the critical unsymmetrical power and loading conditions, 
with appropriate allowance for rigging tolerances.



Sec. 25.409  Tabs.

    (a) Trim tabs. Trim tabs must be designed to withstand loads arising 
from all likely combinations of tab setting, primary control position, 
and airplane speed (obtainable without exceeding the flight load 
conditions prescribed for the airplane as a whole), when the effect of 
the tab is opposed by pilot effort forces up to those specified in Sec. 
25.397(b).
    (b) Balancing tabs. Balancing tabs must be designed for deflections 
consistent with the primary control surface loading conditions.
    (c) Servo tabs. Servo tabs must be designed for deflections 
consistent with

[[Page 416]]

the primary control surface loading conditions obtainable within the 
pilot maneuvering effort, considering possible opposition from the trim 
tabs.



Sec. 25.415  Ground gust conditions.

    (a) The control system must be designed as follows for control 
surface loads due to ground gusts and taxiing downwind:
    (1) The control system between the stops nearest the surfaces and 
the cockpit controls must be designed for loads corresponding to the 
limit hinge moments H of paragraph (a)(2) of this section. These loads 
need not exceed--
    (i) The loads corresponding to the maximum pilot loads in Sec. 
25.397(c) for each pilot alone; or
    (ii) 0.75 times these maximum loads for each pilot when the pilot 
forces are applied in the same direction.
    (2) The control system stops nearest the surfaces, the control 
system locks, and the parts of the systems (if any) between these stops 
and locks and the control surface horns, must be designed for limit 
hinge moments H, in foot pounds, obtained from the formula, 
H=.0034KV\2\cS, where--

V=65 (wind speed in knots)
K=limit hinge moment factor for ground gusts derived in paragraph (b) of 
          this section.
c=mean chord of the control surface aft of the hinge line (ft);
S=area of the control surface aft of the hinge line (sq ft);

    (b) The limit hinge moment factor K for ground gusts must be derived 
as follows:

------------------------------------------------------------------------
               Surface                     K       Position of controls
------------------------------------------------------------------------
(a) Aileron..........................      0.75  Control column locked
                                                  or lashed in mid-
                                                  position.
(b) ......do.........................     \1\ 1  Ailerons at full throw.
                                         
Sec. 25.415  Ground gust conditions.

    (a) The flight control systems and surfaces must be designed for the 
limit loads generated when the airplane is subjected to a horizontal 65-
knot ground gust from any direction while taxiing and while parked. For 
airplanes equipped with control system gust locks, the taxiing condition 
must be evaluated with the controls locked and unlocked, and the parked 
condition must be evaluated with the controls locked.
    (b) The control system and surface loads due to ground gust may be 
assumed to be static loads, and the hinge moments H must be computed 
from the formula:
H = K (1/2) [rho]o V\2\ c S

Where--

K = hinge moment factor for ground gusts derived in paragraph (c) of 
          this section;
[rho]o = density of air at sea level;
V = 65 knots relative to the aircraft;
S = area of the control surface aft of the hinge line;
c = mean aerodynamic chord of the control surface aft of the hinge line.

    (c) The hinge moment factor K for ground gusts must be taken from 
the following table:

------------------------------------------------------------------------
               Surface                    K       Position of  controls
------------------------------------------------------------------------
(1) Aileron.........................       0.75  Control column locked
                                                  or lashed in mid-
                                                  position.
(2) Aileron.........................  * 
Sec. 25.427  Unsymmetrical loads.

    (a) In designing the airplane for lateral gust, yaw maneuver and 
roll maneuver conditions, account must be taken of unsymmetrical loads 
on the empennage arising from effects such as slipstream and aerodynamic 
interference with the wing, vertical fin and other aerodynamic surfaces.
    (b) The horizontal tail must be assumed to be subjected to 
unsymmetrical loading conditions determined as follows:
    (1) 100 percent of the maximum loading from the symmetrical maneuver 
conditions of Sec. 25.331 and the vertical gust conditions of Sec. 
25.341(a) acting separately on the surface on one side of the plane of 
symmetry; and
    (2) 80 percent of these loadings acting on the other side.
    (c) For empennage arrangements where the horizontal tail surfaces 
have dihedral angles greater than plus or minus 10 degrees, or are 
supported by the vertical tail surfaces, the surfaces and the supporting 
structure must be designed for gust velocities specified in Sec. 
25.341(a) acting in any orientation at right angles to the flight path.
    (d) Unsymmetrical loading on the empennage arising from buffet 
conditions of Sec. 25.305(e) must be taken into account.

[Doc. No. 27902, 61 FR 5222, Feb. 9, 1996]



Sec. 25.445  Auxiliary aerodynamic surfaces.

    (a) When significant, the aerodynamic influence between auxiliary 
aerodynamic surfaces, such as outboard fins and winglets, and their 
supporting aerodynamic surfaces, must be taken into account for all 
loading conditions including pitch, roll, and yaw maneuvers, and gusts 
as specified in Sec. 25.341(a) acting at any orientation at right 
angles to the flight path.
    (b) To provide for unsymmetrical loading when outboard fins extend 
above and below the horizontal surface, the critical vertical surface 
loading (load per unit area) determined under Sec. 25.391 must also be 
applied as follows:
    (1) 100 percent to the area of the vertical surfaces above (or 
below) the horizontal surface.
    (2) 80 percent to the area below (or above) the horizontal surface.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-86, 
61 FR 5222, Feb. 9, 1996]



Sec. 25.457  Wing flaps.

    Wing flaps, their operating mechanisms, and their supporting 
structures must be designed for critical loads occurring in the 
conditions prescribed in Sec. 25.345, accounting for the loads 
occurring during transition from one flap position and airspeed to 
another.



Sec. 25.459  Special devices.

    The loading for special devices using aerodynamic surfaces (such as 
slots, slats and spoilers) must be determined from test data.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29776, July 20, 1990]

[[Page 418]]

                              Ground Loads



Sec. 25.471  General.

    (a) Loads and equilibrium. For limit ground loads--
    (1) Limit ground loads obtained under this subpart are considered to 
be external forces applied to the airplane structure; and
    (2) In each specified ground load condition, the external loads must 
be placed in equilibrium with the linear and angular inertia loads in a 
rational or conservative manner.
    (b) Critical centers of gravity. The critical centers of gravity 
within the range for which certification is requested must be selected 
so that the maximum design loads are obtained in each landing gear 
element. Fore and aft, vertical, and lateral airplane centers of gravity 
must be considered. Lateral displacements of the c.g. from the airplane 
centerline which would result in main gear loads not greater than 103 
percent of the critical design load for symmetrical loading conditions 
may be selected without considering the effects of these lateral c.g. 
displacements on the loading of the main gear elements, or on the 
airplane structure provided--
    (1) The lateral displacement of the c.g. results from random 
passenger or cargo disposition within the fuselage or from random 
unsymmetrical fuel loading or fuel usage; and
    (2) Appropriate loading instructions for random disposable loads are 
included under the provisions of Sec. 25.1583(c)(1) to ensure that the 
lateral displacement of the center of gravity is maintained within these 
limits.
    (c) Landing gear dimension data. Figure 1 of appendix A contains the 
basic landing gear dimension data.

[Amdt. 25-23, 35 FR 5673, Apr. 8, 1970]



Sec. 25.473  Landing load conditions and assumptions.

    (a) For the landing conditions specified in Sec. 25.479 to Sec. 
25.485 the airplane is assumed to contact the ground--
    (1) In the attitudes defined in Sec. 25.479 and Sec. 25.481;
    (2) With a limit descent velocity of 10 fps at the design landing 
weight (the maximum weight for landing conditions at maximum descent 
velocity); and
    (3) With a limit descent velocity of 6 fps at the design take-off 
weight (the maximum weight for landing conditions at a reduced descent 
velocity).
    (4) The prescribed descent velocities may be modified if it is shown 
that the airplane has design features that make it impossible to develop 
these velocities.
    (b) Airplane lift, not exceeding airplane weight, may be assumed 
unless the presence of systems or procedures significantly affects the 
lift.
    (c) The method of analysis of airplane and landing gear loads must 
take into account at least the following elements:
    (1) Landing gear dynamic characteristics.
    (2) Spin-up and springback.
    (3) Rigid body response.
    (4) Structural dynamic response of the airframe, if significant.
    (d) The landing gear dynamic characteristics must be validated by 
tests as defined in Sec. 25.723(a).
    (e) The coefficient of friction between the tires and the ground may 
be established by considering the effects of skidding velocity and tire 
pressure. However, this coefficient of friction need not be more than 
0.8.

[Amdt. 25-91, 62 FR 40705, July 29, 1997; Amdt. 25-91, 62 FR 45481, Aug. 
27, 1997; Amdt. 25-103, 66 FR 27394, May 16, 2001]



Sec. 25.477  Landing gear arrangement.

    Sections 25.479 through 25.485 apply to airplanes with conventional 
arrangements of main and nose gears, or main and tail gears, when normal 
operating techniques are used.



Sec. 25.479  Level landing conditions.

    (a) In the level attitude, the airplane is assumed to contact the 
ground at forward velocity components, ranging from VL1 to 
1.25 VL2 parallel to the ground under the conditions 
prescribed in Sec. 25.473 with--
    (1) VL1 equal to VS0 (TAS) at the appropriate 
landing weight and in standard sea level conditions; and
    (2) VL2 equal to VS0 (TAS) at the appropriate 
landing weight and altitudes in a hot day temperature of 41 degrees F. 
above standard.

[[Page 419]]

    (3) The effects of increased contact speed must be investigated if 
approval of downwind landings exceeding 10 knots is requested.
    (b) For the level landing attitude for airplanes with tail wheels, 
the conditions specified in this section must be investigated with the 
airplane horizontal reference line horizontal in accordance with Figure 
2 of Appendix A of this part.
    (c) For the level landing attitude for airplanes with nose wheels, 
shown in Figure 2 of Appendix A of this part, the conditions specified 
in this section must be investigated assuming the following attitudes:
    (1) An attitude in which the main wheels are assumed to contact the 
ground with the nose wheel just clear of the ground; and
    (2) If reasonably attainable at the specified descent and forward 
velocities, an attitude in which the nose and main wheels are assumed to 
contact the ground simultaneously.
    (d) In addition to the loading conditions prescribed in paragraph 
(a) of this section, but with maximum vertical ground reactions 
calculated from paragraph (a), the following apply:
    (1) The landing gear and directly affected attaching structure must 
be designed for the maximum vertical ground reaction combined with an 
aft acting drag component of not less than 25% of this maximum vertical 
ground reaction.
    (2) The most severe combination of loads that are likely to arise 
during a lateral drift landing must be taken into account. In absence of 
a more rational analysis of this condition, the following must be 
investigated:
    (i) A vertical load equal to 75% of the maximum ground reaction of 
Sec. 25.473 must be considered in combination with a drag and side load 
of 40% and 25% respectively of that vertical load.
    (ii) The shock absorber and tire deflections must be assumed to be 
75% of the deflection corresponding to the maximum ground reaction of 
Sec. 25.473(a)(2). This load case need not be considered in combination 
with flat tires.
    (3) The combination of vertical and drag components is considered to 
be acting at the wheel axle centerline.

[Amdt. 25-91, 62 FR 40705, July 29, 1997; Amdt. 25-91, 62 FR 45481, Aug. 
27, 1997]



Sec. 25.481  Tail-down landing conditions.

    (a) In the tail-down attitude, the airplane is assumed to contact 
the ground at forward velocity components, ranging from VL1 
to VL2 parallel to the ground under the conditions prescribed 
in Sec. 25.473 with--
    (1) VL1 equal to VS0 (TAS) at the appropriate 
landing weight and in standard sea level conditions; and
    (2) VL2 equal to VS0 (TAS) at the appropriate 
landing weight and altitudes in a hot day temperature of 41 degrees F. 
above standard.
    (3) The combination of vertical and drag components considered to be 
acting at the main wheel axle centerline.
    (b) For the tail-down landing condition for airplanes with tail 
wheels, the main and tail wheels are assumed to contact the ground 
simultaneously, in accordance with figure 3 of appendix A. Ground 
reaction conditions on the tail wheel are assumed to act--
    (1) Vertically; and
    (2) Up and aft through the axle at 45 degrees to the ground line.
    (c) For the tail-down landing condition for airplanes with nose 
wheels, the airplane is assumed to be at an attitude corresponding to 
either the stalling angle or the maximum angle allowing clearance with 
the ground by each part of the airplane other than the main wheels, in 
accordance with figure 3 of appendix A, whichever is less.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-91, 
62 FR 40705, July 29, 1997; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998]



Sec. 25.483  One-gear landing conditions.

    For the one-gear landing conditions, the airplane is assumed to be 
in the level attitude and to contact the ground on one main landing 
gear, in accordance with Figure 4 of Appendix A of this part. In this 
attitude--
    (a) The ground reactions must be the same as those obtained on that 
side under Sec. 25.479(d)(1), and

[[Page 420]]

    (b) Each unbalanced external load must be reacted by airplane 
inertia in a rational or conservative manner.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-91, 
62 FR 40705, July 29, 1997]



Sec. 25.485  Side load conditions.

    In addition to Sec. 25.479(d)(2) the following conditions must be 
considered:
    (a) For the side load condition, the airplane is assumed to be in 
the level attitude with only the main wheels contacting the ground, in 
accordance with figure 5 of appendix A.
    (b) Side loads of 0.8 of the vertical reaction (on one side) acting 
inward and 0.6 of the vertical reaction (on the other side) acting 
outward must be combined with one-half of the maximum vertical ground 
reactions obtained in the level landing conditions. These loads are 
assumed to be applied at the ground contact point and to be resisted by 
the inertia of the airplane. The drag loads may be assumed to be zero.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-91, 
62 FR 40705, July 29, 1997]



Sec. 25.487  Rebound landing condition.

    (a) The landing gear and its supporting structure must be 
investigated for the loads occurring during rebound of the airplane from 
the landing surface.
    (b) With the landing gear fully extended and not in contact with the 
ground, a load factor of 20.0 must act on the unsprung weights of the 
landing gear. This load factor must act in the direction of motion of 
the unsprung weights as they reach their limiting positions in extending 
with relation to the sprung parts of the landing gear.



Sec. 25.489  Ground handling conditions.

    Unless otherwise prescribed, the landing gear and airplane structure 
must be investigated for the conditions in Sec. Sec. 25.491 through 
25.509 with the airplane at the design ramp weight (the maximum weight 
for ground handling conditions). No wing lift may be considered. The 
shock absorbers and tires may be assumed to be in their static position.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970]



Sec. 25.491  Taxi, takeoff and landing roll.

    Within the range of appropriate ground speeds and approved weights, 
the airplane structure and landing gear are assumed to be subjected to 
loads not less than those obtained when the aircraft is operating over 
the roughest ground that may reasonably be expected in normal operation.

[Amdt. 25-91, 62 FR 40705, July 29, 1997]



Sec. 25.493  Braked roll conditions.

    (a) An airplane with a tail wheel is assumed to be in the level 
attitude with the load on the main wheels, in accordance with figure 6 
of appendix A. The limit vertical load factor is 1.2 at the design 
landing weight and 1.0 at the design ramp weight. A drag reaction equal 
to the vertical reaction multiplied by a coefficient of friction of 0.8, 
must be combined with the vertical ground reaction and applied at the 
ground contact point.
    (b) For an airplane with a nose wheel the limit vertical load factor 
is 1.2 at the design landing weight, and 1.0 at the design ramp weight. 
A drag reaction equal to the vertical reaction, multiplied by a 
coefficient of friction of 0.8, must be combined with the vertical 
reaction and applied at the ground contact point of each wheel with 
brakes. The following two attitudes, in accordance with figure 6 of 
appendix A, must be considered:
    (1) The level attitude with the wheels contacting the ground and the 
loads distributed between the main and nose gear. Zero pitching 
acceleration is assumed.
    (2) The level attitude with only the main gear contacting the ground 
and with the pitching moment resisted by angular acceleration.
    (c) A drag reaction lower than that prescribed in this section may 
be used if it is substantiated that an effective drag force of 0.8 times 
the vertical reaction cannot be attained under any likely loading 
condition.

[[Page 421]]

    (d) An airplane equipped with a nose gear must be designed to 
withstand the loads arising from the dynamic pitching motion of the 
airplane due to sudden application of maximum braking force. The 
airplane is considered to be at design takeoff weight with the nose and 
main gears in contact with the ground, and with a steady-state vertical 
load factor of 1.0. The steady-state nose gear reaction must be combined 
with the maximum incremental nose gear vertical reaction caused by the 
sudden application of maximum braking force as described in paragraphs 
(b) and (c) of this section.
    (e) In the absence of a more rational analysis, the nose gear 
vertical reaction prescribed in paragraph (d) of this section must be 
calculated according to the following formula:
[GRAPHIC] [TIFF OMITTED] TR27MY98.017

Where:

VN=Nose gear vertical reaction.
WT=Design takeoff weight.
A=Horizontal distance between the c.g. of the airplane and the nose 
          wheel.
B=Horizontal distance between the c.g. of the airplane and the line 
          joining the centers of the main wheels.
E=Vertical height of the c.g. of the airplane above the ground in the 
          1.0 g static condition.
[mu]=Coefficient of friction of 0.80.
f=Dynamic response factor; 2.0 is to be used unless a lower factor is 
          substantiated. In the absence of other information, the 
          dynamic response factor f may be defined by the equation:
          [GRAPHIC] [TIFF OMITTED] TR27MY98.018
          
Where:

[xi] is the effective critical damping ratio of the rigid body pitching 
          mode about the main landing gear effective ground contact 
          point.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970; Amdt. 25-97, 63 FR 29072, May 27, 1998]



Sec. 25.495  Turning.

    In the static position, in accordance with figure 7 of appendix A, 
the airplane is assumed to execute a steady turn by nose gear steering, 
or by application of sufficient differential power, so that the limit 
load factors applied at the center of gravity are 1.0 vertically and 0.5 
laterally. The side ground reaction of each wheel must be 0.5 of the 
vertical reaction.



Sec. 25.497  Tail-wheel yawing.

    (a) A vertical ground reaction equal to the static load on the tail 
wheel, in combination with a side component of equal magnitude, is 
assumed.
    (b) If there is a swivel, the tail wheel is assumed to be swiveled 
90[deg] to the airplane longitudinal axis with the resultant load 
passing through the axle.
    (c) If there is a lock, steering device, or shimmy damper the tail 
wheel is also assumed to be in the trailing position with the side load 
acting at the ground contact point.



Sec. 25.499  Nose-wheel yaw and steering.

    (a) A vertical load factor of 1.0 at the airplane center of gravity, 
and a side component at the nose wheel ground contact equal to 0.8 of 
the vertical ground reaction at that point are assumed.
    (b) With the airplane assumed to be in static equilibrium with the 
loads resulting from the use of brakes on one side of the main landing 
gear, the nose gear, its attaching structure, and the fuselage structure 
forward of the center of gravity must be designed for the following 
loads:
    (1) A vertical load factor at the center of gravity of 1.0.
    (2) A forward acting load at the airplane center of gravity of 0.8 
times the vertical load on one main gear.
    (3) Side and vertical loads at the ground contact point on the nose 
gear that are required for static equilibrium.
    (4) A side load factor at the airplane center of gravity of zero.
    (c) If the loads prescribed in paragraph (b) of this section result 
in a nose gear side load higher than 0.8 times the vertical nose gear 
load, the design nose gear side load may be limited to 0.8 times the 
vertical load, with unbalanced yawing moments assumed to be resisted by 
airplane inertia forces.
    (d) For other than the nose gear, its attaching structure, and the 
forward

[[Page 422]]

fuselage structure, the loading conditions are those prescribed in 
paragraph (b) of this section, except that--
    (1) A lower drag reaction may be used if an effective drag force of 
0.8 times the vertical reaction cannot be reached under any likely 
loading condition; and
    (2) The forward acting load at the center of gravity need not exceed 
the maximum drag reaction on one main gear, determined in accordance 
with Sec. 25.493(b).
    (e) With the airplane at design ramp weight, and the nose gear in 
any steerable position, the combined application of full normal steering 
torque and vertical force equal to 1.33 times the maximum static 
reaction on the nose gear must be considered in designing the nose gear, 
its attaching structure, and the forward fuselage structure.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 
25-91, 62 FR 40705, July 29, 1997]



Sec. 25.503  Pivoting.

    (a) The airplane is assumed to pivot about one side of the main gear 
with the brakes on that side locked. The limit vertical load factor must 
be 1.0 and the coefficient of friction 0.8.
    (b) The airplane is assumed to be in static equilibrium, with the 
loads being applied at the ground contact points, in accordance with 
figure 8 of appendix A.



Sec. 25.507  Reversed braking.

    (a) The airplane must be in a three point static ground attitude. 
Horizontal reactions parallel to the ground and directed forward must be 
applied at the ground contact point of each wheel with brakes. The limit 
loads must be equal to 0.55 times the vertical load at each wheel or to 
the load developed by 1.2 times the nominal maximum static brake torque, 
whichever is less.
    (b) For airplanes with nose wheels, the pitching moment must be 
balanced by rotational inertia.
    (c) For airplanes with tail wheels, the resultant of the ground 
reactions must pass through the center of gravity of the airplane.



Sec. 25.509  Towing loads.

    (a) The towing loads specified in paragraph (d) of this section must 
be considered separately. These loads must be applied at the towing 
fittings and must act parallel to the ground. In addition--
    (1) A vertical load factor equal to 1.0 must be considered acting at 
the center of gravity;
    (2) The shock struts and tires must be in their static positions; 
and
    (3) With WT as the design ramp weight, the towing load, FTOW, is--
    (i) 0.3 WT for WT less than 30,000 pounds;
    (ii) (6WT + 450,000)/70 for WT between 30,000 and 100,000 pounds; 
and
    (iii) 0.15 WT for WT over 100,000 pounds.
    (b) For towing points not on the landing gear but near the plane of 
symmetry of the airplane, the drag and side tow load components 
specified for the auxiliary gear apply. For towing points located 
outboard of the main gear, the drag and side tow load components 
specified for the main gear apply. Where the specified angle of swivel 
cannot be reached, the maximum obtainable angle must be used.
    (c) The towing loads specified in paragraph (d) of this section must 
be reacted as follows:
    (1) The side component of the towing load at the main gear must be 
reacted by a side force at the static ground line of the wheel to which 
the load is applied.
    (2) The towing loads at the auxiliary gear and the drag components 
of the towing loads at the main gear must be reacted as follows:
    (i) A reaction with a maximum value equal to the vertical reaction 
must be applied at the axle of the wheel to which the load is applied. 
Enough airplane inertia to achieve equilibrium must be applied.
    (ii) The loads must be reacted by airplane inertia.
    (d) The prescribed towing loads are as follows:

[[Page 423]]



----------------------------------------------------------------------------------------------------------------
                                                                                     Load
             Tow point                      Position        ----------------------------------------------------
                                                                   Magnitude         No.         Direction
----------------------------------------------------------------------------------------------------------------
Main gear..........................  ......................  0.75 FTOW per main         1  Forward, parallel to
                                                              gear unit.                2   drag axis.
                                                                                        3  Forward, at 30[deg]
                                                                                        4   to drag axis.
                                                                                           Aft, parallel to drag
                                                                                            axis.
                                                                                           Aft, at 30[deg] to
                                                                                            drag axis.
Auxiliary gear.....................  Swiveled forward......  1.0 FTOW.............      5  Forward.
                                                                                        6  Aft.
                                     Swiveled aft..........  ......do.............      7  Forward.
                                                                                        8  Aft.
                                     Swiveled 45[deg] from   0.5 FTOW.............      9  Forward, in plane of
                                      forward.                                         10   wheel.
                                                                                           Aft, in plane of
                                                                                            wheel.
                                     Swiveled 45[deg] from   ......do.............     11  Forward, in plane of
                                      aft.                                             12   wheel.
                                                                                           Aft, in plane of
                                                                                            wheel.
----------------------------------------------------------------------------------------------------------------


[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970]



Sec. 25.511  Ground load: unsymmetrical loads on multiple-wheel units.

    (a) General. Multiple-wheel landing gear units are assumed to be 
subjected to the limit ground loads prescribed in this subpart under 
paragraphs (b) through (f) of this section. In addition--
    (1) A tandem strut gear arrangement is a multiple-wheel unit; and
    (2) In determining the total load on a gear unit with respect to the 
provisions of paragraphs (b) through (f) of this section, the transverse 
shift in the load centroid, due to unsymmetrical load distribution on 
the wheels, may be neglected.
    (b) Distribution of limit loads to wheels; tires inflated. The 
distribution of the limit loads among the wheels of the landing gear 
must be established for each landing, taxiing, and ground handling 
condition, taking into account the effects of the following factors:
    (1) The number of wheels and their physical arrangements. For truck 
type landing gear units, the effects of any seesaw motion of the truck 
during the landing impact must be considered in determining the maximum 
design loads for the fore and aft wheel pairs.
    (2) Any differentials in tire diameters resulting from a combination 
of manufacturing tolerances, tire growth, and tire wear. A maximum tire-
diameter differential equal to \2/3\ of the most unfavorable combination 
of diameter variations that is obtained when taking into account 
manufacturing tolerances, tire growth, and tire wear, may be assumed.
    (3) Any unequal tire inflation pressure, assuming the maximum 
variation to be 5 percent of the nominal tire 
inflation pressure.
    (4) A runway crown of zero and a runway crown having a convex upward 
shape that may be approximated by a slope of 1\1/2\ percent with the 
horizontal. Runway crown effects must be considered with the nose gear 
unit on either slope of the crown.
    (5) The airplane attitude.
    (6) Any structural deflections.
    (c) Deflated tires. The effect of deflated tires on the structure 
must be considered with respect to the loading conditions specified in 
paragraphs (d) through (f) of this section, taking into account the 
physical arrangement of the gear components. In addition--
    (1) The deflation of any one tire for each multiple wheel landing 
gear unit, and the deflation of any two critical tires for each landing 
gear unit using four or more wheels per unit, must be considered; and
    (2) The ground reactions must be applied to the wheels with inflated 
tires except that, for multiple-wheel gear units with more than one 
shock strut, a rational distribution of the ground reactions between the 
deflated and inflated tires, accounting for the differences in shock 
strut extensions resulting from a deflated tire, may be used.
    (d) Landing conditions. For one and for two deflated tires, the 
applied load to each gear unit is assumed to be 60 percent and 50 
percent, respectively, of the limit load applied to each gear for

[[Page 424]]

each of the prescribed landing conditions. However, for the drift 
landing condition of Sec. 25.485, 100 percent of the vertical load must 
be applied.
    (e) Taxiing and ground handling conditions. For one and for two 
deflated tires--
    (1) The applied side or drag load factor, or both factors, at the 
center of gravity must be the most critical value up to 50 percent and 
40 percent, respectively, of the limit side or drag load factors, or 
both factors, corresponding to the most severe condition resulting from 
consideration of the prescribed taxiing and ground handling conditions;
    (2) For the braked roll conditions of Sec. 25.493 (a) and (b)(2), 
the drag loads on each inflated tire may not be less than those at each 
tire for the symmetrical load distribution with no deflated tires;
    (3) The vertical load factor at the center of gravity must be 60 
percent and 50 percent, respectively, of the factor with no deflated 
tires, except that it may not be less than 1g; and
    (4) Pivoting need not be considered.
    (f) Towing conditions. For one and for two deflated tires, the 
towing load, FTOW, must be 60 percent and 50 percent, respectively, of 
the load prescribed.



Sec. 25.519  Jacking and tie-down provisions.

    (a) General. The airplane must be designed to withstand the limit 
load conditions resulting from the static ground load conditions of 
paragraph (b) of this section and, if applicable, paragraph (c) of this 
section at the most critical combinations of airplane weight and center 
of gravity. The maximum allowable load at each jack pad must be 
specified.
    (b) Jacking. The airplane must have provisions for jacking and must 
withstand the following limit loads when the airplane is supported on 
jacks--
    (1) For jacking by the landing gear at the maximum ramp weight of 
the airplane, the airplane structure must be designed for a vertical 
load of 1.33 times the vertical static reaction at each jacking point 
acting singly and in combination with a horizontal load of 0.33 times 
the vertical static reaction applied in any direction.
    (2) For jacking by other airplane structure at maximum approved 
jacking weight:
    (i) The airplane structure must be designed for a vertical load of 
1.33 times the vertical reaction at each jacking point acting singly and 
in combination with a horizontal load of 0.33 times the vertical static 
reaction applied in any direction.
    (ii) The jacking pads and local structure must be designed for a 
vertical load of 2.0 times the vertical static reaction at each jacking 
point, acting singly and in combination with a horizontal load of 0.33 
times the vertical static reaction applied in any direction.
    (c) Tie-down. If tie-down points are provided, the main tie-down 
points and local structure must withstand the limit loads resulting from 
a 65-knot horizontal wind from any direction.

[Doc. No. 26129, 59 FR 22102, Apr. 28, 1994]

                               Water Loads



Sec. 25.521  General.

    (a) Seaplanes must be designed for the water loads developed during 
takeoff and landing, with the seaplane in any attitude likely to occur 
in normal operation, and at the appropriate forward and sinking 
velocities under the most severe sea conditions likely to be 
encountered.
    (b) Unless a more rational analysis of the water loads is made, or 
the standards in ANC-3 are used, Sec. Sec. 25.523 through 25.537 apply.
    (c) The requirements of this section and Sec. Sec. 25.523 through 
25.537 apply also to amphibians.



Sec. 25.523  Design weights and center of gravity positions.

    (a) Design weights. The water load requirements must be met at each 
operating weight up to the design landing weight except that, for the 
takeoff condition prescribed in Sec. 25.531, the design water takeoff 
weight (the maximum weight for water taxi and takeoff run) must be used.
    (b) Center of gravity positions. The critical centers of gravity 
within the limits for which certification is requested must be 
considered to reach

[[Page 425]]

maximum design loads for each part of the seaplane structure.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970]



Sec. 25.525  Application of loads.

    (a) Unless otherwise prescribed, the seaplane as a whole is assumed 
to be subjected to the loads corresponding to the load factors specified 
in Sec. 25.527.
    (b) In applying the loads resulting from the load factors prescribed 
in Sec. 25.527, the loads may be distributed over the hull or main 
float bottom (in order to avoid excessive local shear loads and bending 
moments at the location of water load application) using pressures not 
less than those prescribed in Sec. 25.533(b).
    (c) For twin float seaplanes, each float must be treated as an 
equivalent hull on a fictitious seaplane with a weight equal to one-half 
the weight of the twin float seaplane.
    (d) Except in the takeoff condition of Sec. 25.531, the aerodynamic 
lift on the seaplane during the impact is assumed to be \2/3\ of the 
weight of the seaplane.



Sec. 25.527  Hull and main float load factors.

    (a) Water reaction load factors nW must be computed in the following 
manner:
    (1) For the step landing case
    [GRAPHIC] [TIFF OMITTED] TC28SE91.036
    
    (2) For the bow and stern landing cases
    [GRAPHIC] [TIFF OMITTED] TC28SE91.037
    
    (b) The following values are used:
    (1) nW=water reaction load factor (that is, the water reaction 
divided by seaplane weight).
    (2) C1=empirical seaplane operations factor equal to 
0.012 (except that this factor may not be less than that necessary to 
obtain the minimum value of step load factor of 2.33).
    (3) VS0=seaplane stalling speed in knots with flaps 
extended in the appropriate landing position and with no slipstream 
effect.
    (4) [beta]=angle of dead rise at the longitudinal station at which 
the load factor is being determined in accordance with figure 1 of 
appendix B.
    (5) W= seaplane design landing weight in pounds.
    (6) K1=empirical hull station weighing factor, in 
accordance with figure 2 of appendix B.
    (7) rx=ratio of distance, measured parallel to hull reference axis, 
from the center of gravity of the seaplane to the hull longitudinal 
station at which the load factor is being computed to the radius of 
gyration in pitch of the seaplane, the hull reference axis being a 
straight line, in the plane of symmetry, tangential to the keel at the 
main step.
    (c) For a twin float seaplane, because of the effect of flexibility 
of the attachment of the floats to the seaplane, the factor 
K1 may be reduced at the bow and stern to 0.8 of the value 
shown in figure 2 of appendix B. This reduction applies only to the 
design of the carrythrough and seaplane structure.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970]



Sec. 25.529  Hull and main float landing conditions.

    (a) Symmetrical step, bow, and stern landing. For symmetrical step, 
bow, and stern landings, the limit water reaction load factors are those 
computed under Sec. 25.527. In addition--
    (1) For symmetrical step landings, the resultant water load must be 
applied at the keel, through the center of gravity, and must be directed 
perpendicularly to the keel line;
    (2) For symmetrical bow landings, the resultant water load must be 
applied at the keel, one-fifth of the longitudinal distance from the bow 
to the step, and must be directed perpendicularly to the keel line; and
    (3) For symmetrical stern landings, the resultant water load must be 
applied at the keel, at a point 85 percent of the longitudinal distance 
from the step to the stern post, and must be directed perpendicularly to 
the keel line.

[[Page 426]]

    (b) Unsymmetrical landing for hull and single float seaplanes. 
Unsymmetrical step, bow, and stern landing conditions must be 
investigated. In addition--
    (1) The loading for each condition consists of an upward component 
and a side component equal, respectively, to 0.75 and 0.25 tan [beta] 
times the resultant load in the corresponding symmetrical landing 
condition; and
    (2) The point of application and direction of the upward component 
of the load is the same as that in the symmetrical condition, and the 
point of application of the side component is at the same longitudinal 
station as the upward component but is directed inward perpendicularly 
to the plane of symmetry at a point midway between the keel and chine 
lines.
    (c) Unsymmetrical landing; twin float seaplanes. The unsymmetrical 
loading consists of an upward load at the step of each float of 0.75 and 
a side load of 0.25 tan [beta] at one float times the step landing load 
reached under Sec. 25.527. The side load is directed inboard, 
perpendicularly to the plane of symmetry midway between the keel and 
chine lines of the float, at the same longitudinal station as the upward 
load.



Sec. 25.531  Hull and main float takeoff condition.

    For the wing and its attachment to the hull or main float--
    (a) The aerodynamic wing lift is assumed to be zero; and
    (b) A downward inertia load, corresponding to a load factor computed 
from the following formula, must be applied:
[GRAPHIC] [TIFF OMITTED] TC28SE91.038

where--

n=inertia load factor;
CTO=empirical seaplane operations factor equal to 0.004;
VS1=seaplane stalling speed (knots) at the design takeoff 
          weight with the flaps extended in the appropriate takeoff 
          position;
[beta]=angle of dead rise at the main step (degrees); and
W=design water takeoff weight in pounds.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970]



Sec. 25.533  Hull and main float bottom pressures.

    (a) General. The hull and main float structure, including frames and 
bulkheads, stringers, and bottom plating, must be designed under this 
section.
    (b) Local pressures. For the design of the bottom plating and 
stringers and their attachments to the supporting structure, the 
following pressure distributions must be applied:
    (1) For an unflared bottom, the pressure at the chine is 0.75 times 
the pressure at the keel, and the pressures between the keel and chine 
vary linearly, in accordance with figure 3 of appendix B. The pressure 
at the keel (psi) is computed as follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.039

where--

Pk=pressure (p.s.i.) at the keel;
C2=0.00213;
K2=hull station weighing factor, in accordance with figure 2 
          of appendix B;
VS1=seaplane stalling speed (Knots) at the design water 
          takeoff weight with flaps extended in the appropriate takeoff 
          position; and
[beta]k=angle of dead rise at keel, in accordance with figure 1 of 
          appendix B.

    (2) For a flared bottom, the pressure at the beginning of the flare 
is the same as that for an unflared bottom, and the pressure between the 
chine and the beginning of the flare varies linearly, in accordance with 
figure 3 of appendix B. The pressure distribution is the same as that 
prescribed in paragraph (b)(1) of this section for an unflared bottom 
except that the pressure at the chine is computed as follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.040

where--

Pch=pressure (p.s.i.) at the chine;
C3=0.0016;
K2=hull station weighing factor, in accordance with figure 2 
          of appendix B;

[[Page 427]]

VS1=seaplane stalling speed at the design water takeoff 
          weight with flaps extended in the appropriate takeoff 
          position; and
[beta]=angle of dead rise at appropriate station.


The area over which these pressures are applied must simulate pressures 
occurring during high localized impacts on the hull or float, but need 
not extend over an area that would induce critical stresses in the 
frames or in the overall structure.
    (c) Distributed pressures. For the design of the frames, keel, and 
chine structure, the following pressure distributions apply:
    (1) Symmetrical pressures are computed as follows:
    [GRAPHIC] [TIFF OMITTED] TC28SE91.041
    
where--
P=pressure (p.s.i.);
C4=0.078 C1 (with C1 computed under 
          Sec. 25.527);
K2=hull station weighing factor, determined in accordance 
          with figure 2 of appendix B;
VS0=seaplane stalling speed (Knots) with landing flaps 
          extended in the appropriate position and with no slipstream 
          effect; and
VS0=seaplane stalling speed with landing flaps extended in 
          the appropriate position and with no slipstream effect; and 
          [beta]=angle of dead rise at appropriate station.

    (2) The unsymmetrical pressure distribution consists of the 
pressures prescribed in paragraph (c)(1) of this section on one side of 
the hull or main float centerline and one-half of that pressure on the 
other side of the hull or main float centerline, in accordance with 
figure 3 of appendix B.

These pressures are uniform and must be applied simultaneously over the 
entire hull or main float bottom. The loads obtained must be carried 
into the sidewall structure of the hull proper, but need not be 
transmitted in a fore and aft direction as shear and bending loads.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970]



Sec. 25.535  Auxiliary float loads.

    (a) General. Auxiliary floats and their attachments and supporting 
structures must be designed for the conditions prescribed in this 
section. In the cases specified in paragraphs (b) through (e) of this 
section, the prescribed water loads may be distributed over the float 
bottom to avoid excessive local loads, using bottom pressures not less 
than those prescribed in paragraph (g) of this section.
    (b) Step loading. The resultant water load must be applied in the 
plane of symmetry of the float at a point three-fourths of the distance 
from the bow to the step and must be perpendicular to the keel. The 
resultant limit load is computed as follows, except that the value of L 
need not exceed three times the weight of the displaced water when the 
float is completely submerged:
[GRAPHIC] [TIFF OMITTED] TC28SE91.042

where--

L=limit load (lbs.);
C5=0.0053;
VS0=seaplane stalling speed (knots) with landing flaps 
          extended in the appropriate position and with no slipstream 
          effect;
W=seaplane design landing weight in pounds;
[beta]S=angle of dead rise at a station \3/4\ of the distance from the 
          bow to the step, but need not be less than 15 degrees; and
ry=ratio of the lateral distance between the center of gravity and the 
          plane of symmetry of the float to the radius of gyration in 
          roll.

    (c) Bow loading. The resultant limit load must be applied in the 
plane of symmetry of the float at a point one-fourth of the distance 
from the bow to the step and must be perpendicular to the tangent to the 
keel line at that point. The magnitude of the resultant load is that 
specified in paragraph (b) of this section.
    (d) Unsymmetrical step loading. The resultant water load consists of 
a component equal to 0.75 times the load specified in paragraph (a) of 
this section and a side component equal to 3.25 tan [beta] times the 
load specified in paragraph (b) of this section. The side load must be 
applied perpendicularly to the plane of symmetry of the float at a point 
midway between the keel and the chine.

[[Page 428]]

    (e) Unsymmetrical bow loading. The resultant water load consists of 
a component equal to 0.75 times the load specified in paragraph (b) of 
this section and a side component equal to 0.25 tan [beta] times the 
load specified in paragraph (c) of this section. The side load must be 
applied perpendicularly to the plane of symmetry at a point midway 
between the keel and the chine.
    (f) Immersed float condition. The resultant load must be applied at 
the centroid of the cross section of the float at a point one-third of 
the distance from the bow to the step. The limit load components are as 
follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.043

where--

[rho]=mass density of water (slugs/ft.\2\);
V=volume of float (ft.\2\);
Cx=coefficient of drag force, equal to 0.133;
Cy=coefficient of side force, equal to 0.106;
K=0.8, except that lower values may be used if it is shown that the 
          floats are incapable of submerging at a speed of 0.8 
          VS0 in normal operations;
VS0=seaplane stalling speed (knots) with landing flaps 
          extended in the appropriate position and with no slipstream 
          effect; and
g=acceleration due to gravity (ft./sec.\2\).

    (g) Float bottom pressures. The float bottom pressures must be 
established under Sec. 25.533, except that the value of K2 
in the formulae may be taken as 1.0. The angle of dead rise to be used 
in determining the float bottom pressures is set forth in paragraph (b) 
of this section.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970]



Sec. 25.537  Seawing loads.

    Seawing design loads must be based on applicable test data.

                      Emergency Landing Conditions



Sec. 25.561  General.

    (a) The airplane, although it may be damaged in emergency landing 
conditions on land or water, must be designed as prescribed in this 
section to protect each occupant under those conditions.
    (b) The structure must be designed to give each occupant every 
reasonable chance of escaping serious injury in a minor crash landing 
when--
    (1) Proper use is made of seats, belts, and all other safety design 
provisions;
    (2) The wheels are retracted (where applicable); and
    (3) The occupant experiences the following ultimate inertia forces 
acting separately relative to the surrounding structure:
    (i) Upward, 3.0g
    (ii) Forward, 9.0g
    (iii) Sideward, 3.0g on the airframe; and 4.0g on the seats and 
their attachments.
    (iv) Downward, 6.0g
    (v) Rearward, 1.5g
    (c) For equipment, cargo in the passenger compartments and any other 
large masses, the following apply:
    (1) Except as provided in paragraph (c)(2) of this section, these 
items must be positioned so that if they break loose they will be 
unlikely to:
    (i) Cause direct injury to occupants;
    (ii) Penetrate fuel tanks or lines or cause fire or explosion hazard 
by damage to adjacent systems; or
    (iii) Nullify any of the escape facilities provided for use after an 
emergency landing.
    (2) When such positioning is not practical (e.g. fuselage mounted 
engines or auxiliary power units) each such item of mass shall be 
restrained under all loads up to those specified in paragraph (b)(3) of 
this section. The local attachments for these items should be designed 
to withstand 1.33 times the specified loads if these items are subject 
to severe wear and tear through frequent removal (e.g. quick change 
interior items).
    (d) Seats and items of mass (and their supporting structure) must 
not deform under any loads up to those specified in paragraph (b)(3) of 
this section in any manner that would impede

[[Page 429]]

subsequent rapid evacuation of occupants.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5673, Apr. 8, 1970; Amdt. 25-64, 53 FR 17646, May 17, 1988; Amdt. 
25-91, 62 FR 40706, July 29, 1997]



Sec. 25.562  Emergency landing dynamic conditions.

    (a) The seat and restraint system in the airplane must be designed 
as prescribed in this section to protect each occupant during an 
emergency landing condition when--
    (1) Proper use is made of seats, safety belts, and shoulder 
harnesses provided for in the design; and
    (2) The occupant is exposed to loads resulting from the conditions 
prescribed in this section.
    (b) Each seat type design approved for crew or passenger occupancy 
during takeoff and landing must successfully complete dynamic tests or 
be demonstrated by rational analysis based on dynamic tests of a similar 
type seat, in accordance with each of the following emergency landing 
conditions. The tests must be conducted with an occupant simulated by a 
170-pound anthropomorphic test dummy, as defined by 49 CFR Part 572, 
Subpart B, or its equivalent, sitting in the normal upright position.
    (1) A change in downward vertical velocity ([Delta] v) of not less 
than 35 feet per second, with the airplane's longitudinal axis canted 
downward 30 degrees with respect to the horizontal plane and with the 
wings level. Peak floor deceleration must occur in not more than 0.08 
seconds after impact and must reach a minimum of 14g.
    (2) A change in forward longitudinal velocity ([Delta] v) of not 
less than 44 feet per second, with the airplane's longitudinal axis 
horizontal and yawed 10 degrees either right or left, whichever would 
cause the greatest likelihood of the upper torso restraint system (where 
installed) moving off the occupant's shoulder, and with the wings level. 
Peak floor deceleration must occur in not more than 0.09 seconds after 
impact and must reach a minimum of 16g. Where floor rails or floor 
fittings are used to attach the seating devices to the test fixture, the 
rails or fittings must be misaligned with respect to the adjacent set of 
rails or fittings by at least 10 degrees vertically (i.e., out of 
Parallel) with one rolled 10 degrees.
    (c) The following performance measures must not be exceeded during 
the dynamic tests conducted in accordance with paragraph (b) of this 
section:
    (1) Where upper torso straps are used for crewmembers, tension loads 
in individual straps must not exceed 1,750 pounds. If dual straps are 
used for restraining the upper torso, the total strap tension loads must 
not exceed 2,000 pounds.
    (2) The maximum compressive load measured between the pelvis and the 
lumbar column of the anthropomorphic dummy must not exceed 1,500 pounds.
    (3) The upper torso restraint straps (where installed) must remain 
on the occupant's shoulder during the impact.
    (4) The lap safety belt must remain on the occupant's pelvis during 
the impact.
    (5) Each occupant must be protected from serious head injury under 
the conditions prescribed in paragraph (b) of this section. Where head 
contact with seats or other structure can occur, protection must be 
provided so that the head impact does not exceed a Head Injury Criterion 
(HIC) of 1,000 units. The level of HIC is defined by the equation:
[GRAPHIC] [TIFF OMITTED] TC28SE91.044

Where:

t1 is the initial integration time,
t2 is the final integration time, and
a(t) is the total acceleration vs. time curve for the head strike, and 
          where

[[Page 430]]

(t) is in seconds, and (a) is in units of gravity (g).

    (6) Where leg injuries may result from contact with seats or other 
structure, protection must be provided to prevent axially compressive 
loads exceeding 2,250 pounds in each femur.
    (7) The seat must remain attached at all points of attachment, 
although the structure may have yielded.
    (8) Seats must not yield under the tests specified in paragraphs 
(b)(1) and (b)(2) of this section to the extent they would impede rapid 
evacuation of the airplane occupants.

[Amdt. 25-64, 53 FR 17646, May 17, 1988]



Sec. 25.563  Structural ditching provisions.

    Structural strength considerations of ditching provisions must be in 
accordance with Sec. 25.801(e).

                           Fatigue Evaluation



Sec. 25.571  Damage--tolerance and fatigue evaluation of structure.

    (a) General. An evaluation of the strength, detail design, and 
fabrication must show that catastrophic failure due to fatigue, 
corrosion, manufacturing defects, or accidental damage, will be avoided 
throughout the operational life of the airplane. This evaluation must be 
conducted in accordance with the provisions of paragraphs (b) and (e) of 
this section, except as specified in paragraph (c) of this section, for 
each part of the structure that could contribute to a catastrophic 
failure (such as wing, empennage, control surfaces and their systems, 
the fuselage, engine mounting, landing gear, and their related primary 
attachments). For turbojet powered airplanes, those parts that could 
contribute to a catastrophic failure must also be evaluated under 
paragraph (d) of this section. In addition, the following apply:
    (1) Each evaluation required by this section must include--
    (i) The typical loading spectra, temperatures, and humidities 
expected in service;
    (ii) The identification of principal structural elements and detail 
design points, the failure of which could cause catastrophic failure of 
the airplane; and
    (iii) An analysis, supported by test evidence, of the principal 
structural elements and detail design points identified in paragraph 
(a)(1)(ii) of this section.
    (2) The service history of airplanes of similar structural design, 
taking due account of differences in operating conditions and 
procedures, may be used in the evaluations required by this section.
    (3) Based on the evaluations required by this section, inspections 
or other procedures must be established, as necessary, to prevent 
catastrophic failure, and must be included in the Airworthiness 
Limitations section of the Instructions for Continued Airworthiness 
required by Sec. 25.1529. The limit of validity of the engineering data 
that supports the structural maintenance program (hereafter referred to 
as LOV), stated as a number of total accumulated flight cycles or flight 
hours or both, established by this section must also be included in the 
Airworthiness Limitations section of the Instructions for Continued 
Airworthiness required by Sec. 25.1529. Inspection thresholds for the 
following types of structure must be established based on crack growth 
analyses and/or tests, assuming the structure contains an initial flaw 
of the maximum probable size that could exist as a result of 
manufacturing or service-induced damage:
    (i) Single load path structure, and
    (ii) Multiple load path ``fail-safe'' structure and crack arrest 
``fail-safe'' structure, where it cannot be demonstrated that load path 
failure, partial failure, or crack arrest will be detected and repaired 
during normal maintenance, inspection, or operation of an airplane prior 
to failure of the remaining structure.
    (b) Damage-tolerance evaluation. The evaluation must include a 
determination of the probable locations and modes of damage due to 
fatigue, corrosion, or accidental damage. Repeated load and static 
analyses supported by test evidence and (if available) service 
experience must also be incorporated in the evaluation. Special 
consideration for widespread fatigue damage must be included where the 
design is

[[Page 431]]

such that this type of damage could occur. An LOV must be established 
that corresponds to the period of time, stated as a number of total 
accumulated flight cycles or flight hours or both, during which it is 
demonstrated that widespread fatigue damage will not occur in the 
airplane structure. This demonstration must be by full-scale fatigue 
test evidence. The type certificate may be issued prior to completion of 
full-scale fatigue testing, provided the Administrator has approved a 
plan for completing the required tests. In that case, the Airworthiness 
Limitations section of the Instructions for Continued Airworthiness 
required by Sec. 25.1529 must specify that no airplane may be operated 
beyond a number of cycles equal to \1/2\ the number of cycles 
accumulated on the fatigue test article, until such testing is 
completed. The extent of damage for residual strength evaluation at any 
time within the operational life of the airplane must be consistent with 
the initial detectability and subsequent growth under repeated loads. 
The residual strength evaluation must show that the remaining structure 
is able to withstand loads (considered as static ultimate loads) 
corresponding to the following conditions:
    (1) The limit symmetrical maneuvering conditions specified in Sec. 
25.337 at all speeds up to Vc and in Sec. 25.345.
    (2) The limit gust conditions specified in Sec. 25.341 at the 
specified speeds up to VC and in Sec. 25.345.
    (3) The limit rolling conditions specified in Sec. 25.349 and the 
limit unsymmetrical conditions specified in Sec. Sec. 25.367 and 25.427 
(a) through (c), at speeds up to VC.
    (4) The limit yaw maneuvering conditions specified in Sec. 
25.351(a) at the specified speeds up to VC.
    (5) For pressurized cabins, the following conditions:
    (i) The normal operating differential pressure combined with the 
expected external aerodynamic pressures applied simultaneously with the 
flight loading conditions specified in paragraphs (b)(1) through (4) of 
this section, if they have a significant effect.
    (ii) The maximum value of normal operating differential pressure 
(including the expected external aerodynamic pressures during 1 g level 
flight) multiplied by a factor of 1.15, omitting other loads.
    (6) For landing gear and directly-affected airframe structure, the 
limit ground loading conditions specified in Sec. Sec. 25.473, 25.491, 
and 25.493.

If significant changes in structural stiffness or geometry, or both, 
follow from a structural failure, or partial failure, the effect on 
damage tolerance must be further investigated.
    (c) Fatigue (safe-life) evaluation. Compliance with the damage-
tolerance requirements of paragraph (b) of this section is not required 
if the applicant establishes that their application for particular 
structure is impractical. This structure must be shown by analysis, 
supported by test evidence, to be able to withstand the repeated loads 
of variable magnitude expected during its service life without 
detectable cracks. Appropriate safe-life scatter factors must be 
applied.
    (d) Sonic fatigue strength. It must be shown by analysis, supported 
by test evidence, or by the service history of airplanes of similar 
structural design and sonic excitation environment, that--
    (1) Sonic fatigue cracks are not probable in any part of the flight 
structure subject to sonic excitation; or
    (2) Catastrophic failure caused by sonic cracks is not probable 
assuming that the loads prescribed in paragraph (b) of this section are 
applied to all areas affected by those cracks.
    (e) Damage-tolerance (discrete source) evaluation. The airplane must 
be capable of successfully completing a flight during which likely 
structural damage occurs as a result of--
    (1) Impact with a 4-pound bird when the velocity of the airplane 
relative to the bird along the airplane's flight path is equal to 
Vc at sea level or 0.85Vc at 8,000 feet, whichever 
is more critical;
    (2) Uncontained fan blade impact;
    (3) Uncontained engine failure; or
    (4) Uncontained high energy rotating machinery failure.

The damaged structure must be able to withstand the static loads 
(considered as ultimate loads) which are reasonably expected to occur on 
the flight. Dynamic effects on these static loads

[[Page 432]]

need not be considered. Corrective action to be taken by the pilot 
following the incident, such as limiting maneuvers, avoiding turbulence, 
and reducing speed, must be considered. If significant changes in 
structural stiffness or geometry, or both, follow from a structural 
failure or partial failure, the effect on damage tolerance must be 
further investigated.

[Amdt. 25-45, 43 FR 46242, Oct. 5, 1978, as amended by Amdt. 25-54, 45 
FR 60173, Sept. 11, 1980; Amdt. 25-72, 55 FR 29776, July 20, 1990; Amdt. 
25-86, 61 FR 5222, Feb. 9, 1996; Amdt. 25-96, 63 FR 15714, Mar. 31, 
1998; 63 FR 23338, Apr. 28, 1998; Amdt. 25-132, 75 FR 69781, Nov. 15, 
2010]

                          Lightning Protection



Sec. 25.581  Lightning protection.

    (a) The airplane must be protected against catastrophic effects from 
lightning.
    (b) For metallic components, compliance with paragraph (a) of this 
section may be shown by--
    (1) Bonding the components properly to the airframe; or
    (2) Designing the components so that a strike will not endanger the 
airplane.
    (c) For nonmetallic components, compliance with paragraph (a) of 
this section may be shown by--
    (1) Designing the components to minimize the effect of a strike; or
    (2) Incorporating acceptable means of diverting the resulting 
electrical current so as not to endanger the airplane.

[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]



                    Subpart D_Design and Construction

                                 General



Sec. 25.601  General.

    The airplane may not have design features or details that experience 
has shown to be hazardous or unreliable. The suitability of each 
questionable design detail and part must be established by tests.



Sec. 25.603  Materials.

    The suitability and durability of materials used for parts, the 
failure of which could adversely affect safety, must--
    (a) Be established on the basis of experience or tests;
    (b) Conform to approved specifications (such as industry or military 
specifications, or Technical Standard Orders) that ensure their having 
the strength and other properties assumed in the design data; and
    (c) Take into account the effects of environmental conditions, such 
as temperature and humidity, expected in service.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55466, Dec. 20 1976; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978]



Sec. 25.605  Fabrication methods.

    (a) The methods of fabrication used must produce a consistently 
sound structure. If a fabrication process (such as gluing, spot welding, 
or heat treating) requires close control to reach this objective, the 
process must be performed under an approved process specification.
    (b) Each new aircraft fabrication method must be substantiated by a 
test program.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 
43 FR 50595, Oct. 30, 1978]



Sec. 25.607  Fasteners.

    (a) Each removable bolt, screw, nut, pin, or other removable 
fastener must incorporate two separate locking devices if--
    (1) Its loss could preclude continued flight and landing within the 
design limitations of the airplane using normal pilot skill and 
strength; or
    (2) Its loss could result in reduction in pitch, yaw, or roll 
control capability or response below that required by Subpart B of this 
chapter.
    (b) The fasteners specified in paragraph (a) of this section and 
their locking devices may not be adversely affected by the environmental 
conditions associated with the particular installation.
    (c) No self-locking nut may be used on any bolt subject to rotation 
in operation unless a nonfriction locking device is used in addition to 
the self-locking device.

[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]

[[Page 433]]



Sec. 25.609  Protection of structure.

    Each part of the structure must--
    (a) Be suitably protected against deterioration or loss of strength 
in service due to any cause, including--
    (1) Weathering;
    (2) Corrosion; and
    (3) Abrasion; and
    (b) Have provisions for ventilation and drainage where necessary for 
protection.



Sec. 25.611  Accessibility provisions.

    (a)Means must be provided to allow inspection (including inspection 
of principal structural elements and control systems), replacement of 
parts normally requiring replacement, adjustment, and lubrication as 
necessary for continued airworthiness. The inspection means for each 
item must be practicable for the inspection interval for the item. 
Nondestructive inspection aids may be used to inspect structural 
elements where it is impracticable to provide means for direct visual 
inspection if it is shown that the inspection is effective and the 
inspection procedures are specified in the maintenance manual required 
by Sec. 25.1529.
    (b) EWIS must meet the accessibility requirements of Sec. 25.1719.

[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970, as amended by Amdt. 25-123, 72 
FR 63404, Nov. 8, 2007]



Sec. 25.613  Material strength properties and material design values.

    (a) Material strength properties must be based on enough tests of 
material meeting approved specifications to establish design values on a 
statistical basis.
    (b) Material design values must be chosen to minimize the 
probability of structural failures due to material variability. Except 
as provided in paragraphs (e) and (f) of this section, compliance must 
be shown by selecting material design values which assure material 
strength with the following probability:
    (1) Where applied loads are eventually distributed through a single 
member within an assembly, the failure of which would result in loss of 
structural integrity of the component, 99 percent probability with 95 
percent confidence.
    (2) For redundant structure, in which the failure of individual 
elements would result in applied loads being safely distributed to other 
load carrying members, 90 percent probability with 95 percent 
confidence.
    (c) The effects of environmental conditions, such as temperature and 
moisture, on material design values used in an essential component or 
structure must be considered where these effects are significant within 
the airplane operating envelope.
    (d) [Reserved]
    (e) Greater material design values may be used if a ``premium 
selection'' of the material is made in which a specimen of each 
individual item is tested before use to determine that the actual 
strength properties of that particular item will equal or exceed those 
used in design.
    (f) Other material design values may be used if approved by the 
Administrator.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 
43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55 FR 29776, July 20, 1990; 
Amdt. 25-112, 68 FR 46431, Aug. 5, 2003]



Sec. 25.619  Special factors.

    The factor of safety prescribed in Sec. 25.303 must be multiplied 
by the highest pertinent special factor of safety prescribed in 
Sec. Sec. 25.621 through 25.625 for each part of the structure whose 
strength is--
    (a) Uncertain;
    (b) Likely to deteriorate in service before normal replacement; or
    (c) Subject to appreciable variability because of uncertainties in 
manufacturing processes or inspection methods.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5674, Apr. 8, 1970]



Sec. 25.621  Casting factors.

    (a) General. For castings used in structural applications, the 
factors, tests, and inspections specified in paragraphs (b) through (d) 
of this section must be applied in addition to those necessary to 
establish foundry quality control. The inspections must meet approved 
specifications. Paragraphs (c) and (d) of this section apply to any 
structural castings, except castings

[[Page 434]]

that are pressure tested as parts of hydraulic or other fluid systems 
and do not support structural loads.
    (b) Bearing stresses and surfaces. The casting factors specified in 
paragraphs (c) and (d) of this section--
    (1) Need not exceed 1.25 with respect to bearing stresses regardless 
of the method of inspection used; and
    (2) Need not be used with respect to the bearing surfaces of a part 
whose bearing factor is larger than the applicable casting factor.
    (c) Critical castings. Each casting whose failure could preclude 
continued safe flight and landing of the airplane or could result in 
serious injury to occupants is a critical casting. Each critical casting 
must have a factor associated with it for showing compliance with 
strength and deformation requirements of Sec. 25.305, and must comply 
with the following criteria associated with that factor:
    (1) A casting factor of 1.0 or greater may be used, provided that--
    (i) It is demonstrated, in the form of process qualification, proof 
of product, and process monitoring that, for each casting design and 
part number, the castings produced by each foundry and process 
combination have coefficients of variation of the material properties 
that are equivalent to those of wrought alloy products of similar 
composition. Process monitoring must include testing of coupons cut from 
the prolongations of each casting (or each set of castings, if produced 
from a single pour into a single mold in a runner system) and, on a 
sampling basis, coupons cut from critical areas of production castings. 
The acceptance criteria for the process monitoring inspections and tests 
must be established and included in the process specifications to ensure 
the properties of the production castings are controlled to within 
levels used in design.
    (ii) Each casting receives:
    (A) Inspection of 100 percent of its surface, using visual 
inspection and liquid penetrant or equivalent inspection methods; and
    (B) Inspection of structurally significant internal areas and areas 
where defects are likely to occur, using radiographic or equivalent 
inspection methods.
    (iii) One casting undergoes a static test and is shown to meet the 
strength and deformation requirements of Sec. 25.305(a) and (b).
    (2) A casting factor of 1.25 or greater may be used, providedthat--
    (i) Each casting receives:
    (A) Inspection of 100 percent of its surface, using visual 
inspection and liquid penetrant or equivalent inspection methods; and
    (B) Inspection of structurally significant internal areas and areas 
where defects are likely to occur, using radiographic or equivalent 
inspection methods.
    (ii) Three castings undergo static tests and are shown to meet:
    (A) The strength requirements of Sec. 25.305(b) at an ultimate load 
corresponding to a casting factor of 1.25; and
    (B) The deformation requirements of Sec. 25.305(a) at a load of 
1.15 times the limit load.
    (3) A casting factor of 1.50 or greater may be used, provided that--
    (i) Each casting receives:
    (A) Inspection of 100 percent of its surface, using visual 
inspection and liquid penetrant or equivalent inspection methods; and
    (B) Inspection of structurally significant internal areas and areas 
where defects are likely to occur, using radiographic or equivalent 
inspection methods.
    (ii) One casting undergoes a static test and is shown to meet:
    (A) The strength requirements of Sec. 25.305(b) at an ultimate load 
corresponding to a casting factor of 1.50; and
    (B) The deformation requirements of Sec. 25.305(a) at a load of 
1.15 times the limit load.
    (d) Non-critical castings. For each casting other than critical 
castings, as specified in paragraph (c) of this section, the following 
apply:
    (1) A casting factor of 1.0 or greater may be used, provided that 
the requirements of (c)(1) of this section are met, or all of the 
following conditions are met:
    (i) Castings are manufactured to approved specifications that 
specify the minimum mechanical properties of the material in the casting 
and provides

[[Page 435]]

for demonstration of these properties by testing of coupons cut from the 
castings on a sampling basis.
    (ii) Each casting receives:
    (A) Inspection of 100 percent of its surface, using visual 
inspection and liquid penetrant or equivalent inspection methods; and
    (B) Inspection of structurally significant internal areas and areas 
where defects are likely to occur, using radiographic or equivalent 
inspection methods.
    (iii) Three sample castings undergo static tests and are shown to 
meet the strength and deformation requirements of Sec. 25.305(a) and 
(b).
    (2) A casting factor of 1.25 or greater may be used, provided that 
each casting receives:
    (i) Inspection of 100 percent of its surface, using visual 
inspection and liquid penetrant or equivalent inspection methods; and
    (ii) Inspection of structurally significant internal areas and areas 
where defects are likely to occur, using radiographic or equivalent 
inspection methods.
    (3) A casting factor of 1.5 or greater may be used, provided that 
each casting receives inspection of 100 percent of its surface using 
visual inspection and liquid penetrant or equivalent inspection methods.
    (4) A casting factor of 2.0 or greater may be used, provided that 
each casting receives inspection of 100 percent of its surface using 
visual inspection methods.
    (5) The number of castings per production batch to be inspected by 
non-visual methods in accordance with paragraphs (d)(2) and (3) of this 
section may be reduced when an approved quality control procedure is 
established.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. No.25-
139, 79 FR 59429, Oct. 2, 2014]



Sec. 25.623  Bearing factors.

    (a) Except as provided in paragraph (b) of this section, each part 
that has clearance (free fit), and that is subject to pounding or 
vibration, must have a bearing factor large enough to provide for the 
effects of normal relative motion.
    (b) No bearing factor need be used for a part for which any larger 
special factor is prescribed.



Sec. 25.625  Fitting factors.

    For each fitting (a part or terminal used to join one structural 
member to another), the following apply:
    (a) For each fitting whose strength is not proven by limit and 
ultimate load tests in which actual stress conditions are simulated in 
the fitting and surrounding structures, a fitting factor of at least 
1.15 must be applied to each part of--
    (1) The fitting;
    (2) The means of attachment; and
    (3) The bearing on the joined members.
    (b) No fitting factor need be used--
    (1) For joints made under approved practices and based on 
comprehensive test data (such as continuous joints in metal plating, 
welded joints, and scarf joints in wood); or
    (2) With respect to any bearing surface for which a larger special 
factor is used.
    (c) For each integral fitting, the part must be treated as a fitting 
up to the point at which the section properties become typical of the 
member.
    (d) For each seat, berth, safety belt, and harness, the fitting 
factor specified in Sec. 25.785(f)(3) applies.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5674, Apr. 8, 1970; Amdt. 25-72, 55 FR 29776, July 20, 1990]



Sec. 25.629  Aeroelastic stability requirements.

    (a) General. The aeroelastic stability evaluations required under 
this section include flutter, divergence, control reversal and any undue 
loss of stability and control as a result of structural deformation. The 
aeroelastic evaluation must include whirl modes associated with any 
propeller or rotating device that contributes significant dynamic 
forces. Compliance with this section must be shown by analyses, wind 
tunnel tests, ground vibration tests, flight tests, or other means found 
necessary by the Administrator.
    (b) Aeroelastic stability envelopes. The airplane must be designed 
to be free

[[Page 436]]

from aeroelastic instability for all configurations and design 
conditions within the aeroelastic stability envelopes as follows:
    (1) For normal conditions without failures, malfunctions, or adverse 
conditions, all combinations of altitudes and speeds encompassed by the 
VD/MD versus altitude envelope enlarged at all 
points by an increase of 15 percent in equivalent airspeed at both 
constant Mach number and constant altitude. In addition, a proper margin 
of stability must exist at all speeds up to VD/MD 
and, there must be no large and rapid reduction in stability as 
VD/MD is approached. The enlarged envelope may be 
limited to Mach 1.0 when MD is less than 1.0 at all design 
altitudes, and
    (2) For the conditions described in Sec. 25.629(d) below, for all 
approved altitudes, any airspeed up to the greater airspeed defined by;
    (i) The VD/MD envelope determined by Sec. 
25.335(b); or,
    (ii) An altitude-airspeed envelope defined by a 15 percent increase 
in equivalent airspeed above VC at constant altitude, from 
sea level to the altitude of the intersection of 1.15 VC with 
the extension of the constant cruise Mach number line, MC, 
then a linear variation in equivalent airspeed to MC+.05 at 
the altitude of the lowest VC/MC intersection; 
then, at higher altitudes, up to the maximum flight altitude, the 
boundary defined by a .05 Mach increase in MC at constant 
altitude.
    (c) Balance weights. If concentrated balance weights are used, their 
effectiveness and strength, including supporting structure, must be 
substantiated.
    (d) Failures, malfunctions, and adverse conditions. The failures, 
malfunctions, and adverse conditions which must be considered in showing 
compliance with this section are:
    (1) Any critical fuel loading conditions, not shown to be extremely 
improbable, which may result from mismanagement of fuel.
    (2) Any single failure in any flutter damper system.
    (3) For airplanes not approved for operation in icing conditions, 
the maximum likely ice accumulation expected as a result of an 
inadvertent encounter.
    (4) Failure of any single element of the structure supporting any 
engine, independently mounted propeller shaft, large auxiliary power 
unit, or large externally mounted aerodynamic body (such as an external 
fuel tank).
    (5) For airplanes with engines that have propellers or large 
rotating devices capable of significant dynamic forces, any single 
failure of the engine structure that would reduce the rigidity of the 
rotational axis.
    (6) The absence of aerodynamic or gyroscopic forces resulting from 
the most adverse combination of feathered propellers or other rotating 
devices capable of significant dynamic forces. In addition, the effect 
of a single feathered propeller or rotating device must be coupled with 
the failures of paragraphs (d)(4) and (d)(5) of this section.
    (7) Any single propeller or rotating device capable of significant 
dynamic forces rotating at the highest likely overspeed.
    (8) Any damage or failure condition, required or selected for 
investigation by Sec. 25.571. The single structural failures described 
in paragraphs (d)(4) and (d)(5) of this section need not be considered 
in showing compliance with this section if;
    (i) The structural element could not fail due to discrete source 
damage resulting from the conditions described in Sec. 25.571(e), and
    (ii) A damage tolerance investigation in accordance with Sec. 
25.571(b) shows that the maximum extent of damage assumed for the 
purpose of residual strength evaluation does not involve complete 
failure of the structural element.
    (9) Any damage, failure, or malfunction considered under Sec. Sec. 
25.631, 25.671, 25.672, and 25.1309.
    (10) Any other combination of failures, malfunctions, or adverse 
conditions not shown to be extremely improbable.
    (e) Flight flutter testing. Full scale flight flutter tests at 
speeds up to VDF/MDF must be conducted for new 
type designs and for modifications to a type design unless the 
modifications have been shown to have an insignificant effect on the 
aeroelastic stability. These tests must demonstrate that the airplane 
has a proper margin of damping

[[Page 437]]

at all speeds up to VDF/MDF, and that there is no 
large and rapid reduction in damping as VDF/MDF, 
is approached. If a failure, malfunction, or adverse condition is 
simulated during flight test in showing compliance with paragraph (d) of 
this section, the maximum speed investigated need not exceed 
VFC/MFC if it is shown, by correlation of the 
flight test data with other test data or analyses, that the airplane is 
free from any aeroelastic instability at all speeds within the altitude-
airspeed envelope described in paragraph (b)(2) of this section.

[Doc. No. 26007, 57 FR 28949, June 29, 1992]



Sec. 25.631  Bird strike damage.

    The empennage structure must be designed to assure capability of 
continued safe flight and landing of the airplane after impact with an 
8-pound bird when the velocity of the airplane (relative to the bird 
along the airplane's flight path) is equal to VC at sea level, selected 
under Sec. 25.335(a). Compliance with this section by provision of 
redundant structure and protected location of control system elements or 
protective devices such as splitter plates or energy absorbing material 
is acceptable. Where compliance is shown by analysis, tests, or both, 
use of data on airplanes having similar structural design is acceptable.

[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]

                            Control Surfaces



Sec. 25.651  Proof of strength.

    (a) Limit load tests of control surfaces are required. These tests 
must include the horn or fitting to which the control system is 
attached.
    (b) Compliance with the special factors requirements of Sec. Sec. 
25.619 through 25.625 and 25.657 for control surface hinges must be 
shown by analysis or individual load tests.



Sec. 25.655  Installation.

    (a) Movable tail surfaces must be installed so that there is no 
interference between any surfaces when one is held in its extreme 
position and the others are operated through their full angular 
movement.
    (b) If an adjustable stabilizer is used, it must have stops that 
will limit its range of travel to the maximum for which the airplane is 
shown to meet the trim requirements of Sec. 25.161.



Sec. 25.657  Hinges.

    (a) For control surface hinges, including ball, roller, and self-
lubricated bearing hinges, the approved rating of the bearing may not be 
exceeded. For nonstandard bearing hinge configurations, the rating must 
be established on the basis of experience or tests and, in the absence 
of a rational investigation, a factor of safety of not less than 6.67 
must be used with respect to the ultimate bearing strength of the 
softest material used as a bearing.
    (b) Hinges must have enough strength and rigidity for loads parallel 
to the hinge line.

[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]

                             Control Systems



Sec. 25.671  General.

    (a) Each control and control system must operate with the ease, 
smoothness, and positiveness appropriate to its function.
    (b) Each element of each flight control system must be designed, or 
distinctively and permanently marked, to minimize the probability of 
incorrect assembly that could result in the malfunctioning of the 
system.
    (c) The airplane must be shown by analysis, tests, or both, to be 
capable of continued safe flight and landing after any of the following 
failures or jamming in the flight control system and surfaces (including 
trim, lift, drag, and feel systems), within the normal flight envelope, 
without requiring exceptional piloting skill or strength. Probable 
malfunctions must have only minor effects on control system operation 
and must be capable of being readily counteracted by the pilot.
    (1) Any single failure, excluding jamming (for example, 
disconnection or failure of mechanical elements, or structural failure 
of hydraulic components, such as actuators, control spool housing, and 
valves).
    (2) Any combination of failures not shown to be extremely 
improbable, excluding jamming (for example, dual

[[Page 438]]

electrical or hydraulic system failures, or any single failure in 
combination with any probable hydraulic or electrical failure).
    (3) Any jam in a control position normally encountered during 
takeoff, climb, cruise, normal turns, descent, and landing unless the 
jam is shown to be extremely improbable, or can be alleviated. A runaway 
of a flight control to an adverse position and jam must be accounted for 
if such runaway and subsequent jamming is not extremely improbable.
    (d) The airplane must be designed so that it is controllable if all 
engines fail. Compliance with this requirement may be shown by analysis 
where that method has been shown to be reliable.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5674, Apr. 8, 1970]



Sec. 25.672  Stability augmentation and automatic and power-operated systems.

    If the functioning of stability augmentation or other automatic or 
power-operated systems is necessary to show compliance with the flight 
characteristics requirements of this part, such systems must comply with 
Sec. 25.671 and the following:
    (a) A warning which is clearly distinguishable to the pilot under 
expected flight conditions without requiring his attention must be 
provided for any failure in the stability augmentation system or in any 
other automatic or power-operated system which could result in an unsafe 
condition if the pilot were not aware of the failure. Warning systems 
must not activate the control systems.
    (b) The design of the stability augmentation system or of any other 
automatic or power-operated system must permit initial counteraction of 
failures of the type specified in Sec. 25.671(c) without requiring 
exceptional pilot skill or strength, by either the deactivation of the 
system, or a failed portion thereof, or by overriding the failure by 
movement of the flight controls in the normal sense.
    (c) It must be shown that after any single failure of the stability 
augmentation system or any other automatic or power-operated system--
    (1) The airplane is safely controllable when the failure or 
malfunction occurs at any speed or altitude within the approved 
operating limitations that is critical for the type of failure being 
considered;
    (2) The controllability and maneuverability requirements of this 
part are met within a practical operational flight envelope (for 
example, speed, altitude, normal acceleration, and airplane 
configurations) which is described in the Airplane Flight Manual; and
    (3) The trim, stability, and stall characteristics are not impaired 
below a level needed to permit continued safe flight and landing.

[Amdt. 25-23, 35 FR 5675 Apr. 8, 1970]



Sec. 25.675  Stops.

    (a) Each control system must have stops that positively limit the 
range of motion of each movable aerodynamic surface controlled by the 
system.
    (b) Each stop must be located so that wear, slackness, or take-up 
adjustments will not adversely affect the control characteristics of the 
airplane because of a change in the range of surface travel.
    (c) Each stop must be able to withstand any loads corresponding to 
the design conditions for the control system.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55466, Dec. 20, 1976]



Sec. 25.677  Trim systems.

    (a) Trim controls must be designed to prevent inadvertent or abrupt 
operation and to operate in the plane, and with the sense of motion, of 
the airplane.
    (b) There must be means adjacent to the trim control to indicate the 
direction of the control movement relative to the airplane motion. In 
addition, there must be clearly visible means to indicate the position 
of the trim device with respect to the range of adjustment. The 
indicator must be clearly marked with the range within which it has been 
demonstrated that takeoff is safe for all center of gravity positions 
approved for takeoff.

[[Page 439]]

    (c) Trim control systems must be designed to prevent creeping in 
flight. Trim tab controls must be irreversible unless the tab is 
appropriately balanced and shown to be free from flutter.
    (d) If an irreversible tab control system is used, the part from the 
tab to the attachment of the irreversible unit to the airplane structure 
must consist of a rigid connection.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5675, Apr. 8, 1970; Amdt. 25-115, 69 FR 40527, July 2, 2004]



Sec. 25.679  Control system gust locks.

    (a) There must be a device to prevent damage to the control surfaces 
(including tabs), and to the control system, from gusts striking the 
airplane while it is on the ground or water. If the device, when 
engaged, prevents normal operation of the control surfaces by the pilot, 
it must--
    (1) Automatically disengage when the pilot operates the primary 
flight controls in a normal manner; or
    (2) Limit the operation of the airplane so that the pilot receives 
unmistakable warning at the start of takeoff.
    (b) The device must have means to preclude the possibility of it 
becoming inadvertently engaged in flight.



Sec. 25.681  Limit load static tests.

    (a) Compliance with the limit load requirements of this Part must be 
shown by tests in which--
    (1) The direction of the test loads produces the most severe loading 
in the control system; and
    (2) Each fitting, pulley, and bracket used in attaching the system 
to the main structure is included.
    (b) Compliance must be shown (by analyses or individual load tests) 
with the special factor requirements for control system joints subject 
to angular motion.



Sec. 25.683  Operation tests.

    (a) It must be shown by operation tests that when portions of the 
control system subject to pilot effort loads are loaded to 80 percent of 
the limit load specified for the system and the powered portions of the 
control system are loaded to the maximum load expected in normal 
operation, the system is free from--
    (1) Jamming;
    (2) Excessive friction; and
    (3) Excessive deflection.
    (b) It must be shown by analysis and, where necessary, by tests, 
that in the presence of deflections of the airplane structure due to the 
separate application of pitch, roll, and yaw limit maneuver loads, the 
control system, when loaded to obtain these limit loads and operated 
within its operational range of deflections, can be exercised about all 
control axes and remain free from--
    (1) Jamming;
    (2) Excessive friction;
    (3) Disconnection; and
    (4) Any form of permanent damage.
    (c) It must be shown that under vibration loads in the normal flight 
and ground operating conditions, no hazard can result from interference 
or contact with adjacent elements.

[Amdt. No. 25-139, 79 FR 59430, Oct. 2, 2014]



Sec. 25.685  Control system details.

    (a) Each detail of each control system must be designed and 
installed to prevent jamming, chafing, and interference from cargo, 
passengers, loose objects, or the freezing of moisture.
    (b) There must be means in the cockpit to prevent the entry of 
foreign objects into places where they would jam the system.
    (c) There must be means to prevent the slapping of cables or tubes 
against other parts.
    (d) Sections 25.689 and 25.693 apply to cable systems and joints.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55466, Dec. 20, 1976]



Sec. 25.689  Cable systems.

    (a) Each cable, cable fitting, turnbuckle, splice, and pulley must 
be approved. In addition--
    (1) No cable smaller than \1/8\ inch in diameter may be used in the 
aileron, elevator, or rudder systems; and
    (2) Each cable system must be designed so that there will be no 
hazardous change in cable tension throughout the range of travel under 
operating conditions and temperature variations.

[[Page 440]]

    (b) Each kind and size of pulley must correspond to the cable with 
which it is used. Pulleys and sprockets must have closely fitted guards 
to prevent the cables and chains from being displaced or fouled. Each 
pulley must lie in the plane passing through the cable so that the cable 
does not rub against the pulley flange.
    (c) Fairleads must be installed so that they do not cause a change 
in cable direction of more than three degrees.
    (d) Clevis pins subject to load or motion and retained only by 
cotter pins may not be used in the control system.
    (e) Turnbuckles must be attached to parts having angular motion in a 
manner that will positively prevent binding throughout the range of 
travel.
    (f) There must be provisions for visual inspection of fairleads, 
pulleys, terminals, and turnbuckles.



Sec. 25.693  Joints.

    Control system joints (in push-pull systems) that are subject to 
angular motion, except those in ball and roller bearing systems, must 
have a special factor of safety of not less than 3.33 with respect to 
the ultimate bearing strength of the softest material used as a bearing. 
This factor may be reduced to 2.0 for joints in cable control systems. 
For ball or roller bearings, the approved ratings may not be exceeded.

[Amdt. 25-72, 55 FR 29777, July 20, 1990]



Sec. 25.697  Lift and drag devices, controls.

    (a) Each lift device control must be designed so that the pilots can 
place the device in any takeoff, en route, approach, or landing position 
established under Sec. 25.101(d). Lift and drag devices must maintain 
the selected positions, except for movement produced by an automatic 
positioning or load limiting device, without further attention by the 
pilots.
    (b) Each lift and drag device control must be designed and located 
to make inadvertent operation improbable. Lift and drag devices intended 
for ground operation only must have means to prevent the inadvertant 
operation of their controls in flight if that operation could be 
hazardous.
    (c) The rate of motion of the surfaces in response to the operation 
of the control and the characteristics of the automatic positioning or 
load limiting device must give satisfactory flight and performance 
characteristics under steady or changing conditions of airspeed, engine 
power, and airplane attitude.
    (d) The lift device control must be designed to retract the surfaces 
from the fully extended position, during steady flight at maximum 
continuous engine power at any speed below VF +9.0 (knots).

[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970, as amended by Amdt. 25-46, 43 FR 
50595, Oct. 30, 1978; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984]



Sec. 25.699  Lift and drag device indicator.

    (a) There must be means to indicate to the pilots the position of 
each lift or drag device having a separate control in the cockpit to 
adjust its position. In addition, an indication of unsymmetrical 
operation or other malfunction in the lift or drag device systems must 
be provided when such indication is necessary to enable the pilots to 
prevent or counteract an unsafe flight or ground condition, considering 
the effects on flight characteristics and performance.
    (b) There must be means to indicate to the pilots the takeoff, en 
route, approach, and landing lift device positions.
    (c) If any extension of the lift and drag devices beyond the landing 
position is possible, the controls must be clearly marked to identify 
this range of extension.

[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]



Sec. 25.701  Flap and slat interconnection.

    (a) Unless the airplane has safe flight characteristics with the 
flaps or slats retracted on one side and extended on the other, the 
motion of flaps or slats on opposite sides of the plane of symmetry must 
be synchronized by a mechanical interconnection or approved equivalent 
means.
    (b) If a wing flap or slat interconnection or equivalent means is 
used, it must be designed to account for the applicable unsymmetrical 
loads, including those resulting from flight with the

[[Page 441]]

engines on one side of the plane of symmetry inoperative and the 
remaining engines at takeoff power.
    (c) For airplanes with flaps or slats that are not subjected to 
slipstream conditions, the structure must be designed for the loads 
imposed when the wing flaps or slats on one side are carrying the most 
severe load occurring in the prescribed symmetrical conditions and those 
on the other side are carrying not more than 80 percent of that load.
    (d) The interconnection must be designed for the loads resulting 
when interconnected flap or slat surfaces on one side of the plane of 
symmetry are jammed and immovable while the surfaces on the other side 
are free to move and the full power of the surface actuating system is 
applied.

[Amdt. 25-72, 55 FR 29777, July 20, 1990]



Sec. 25.703  Takeoff warning system.

    A takeoff warning system must be installed and must meet the 
following requirements:
    (a) The system must provide to the pilots an aural warning that is 
automatically activated during the initial portion of the takeoff roll 
if the airplane is in a configuration, including any of the following, 
that would not allow a safe takeoff:
    (1) The wing flaps or leading edge devices are not within the 
approved range of takeoff positions.
    (2) Wing spoilers (except lateral control spoilers meeting the 
requirements of Sec. 25.671), speed brakes, or longitudinal trim 
devices are in a position that would not allow a safe takeoff.
    (b) The warning required by paragraph (a) of this section must 
continue until--
    (1) The configuration is changed to allow a safe takeoff;
    (2) Action is taken by the pilot to terminate the takeoff roll;
    (3) The airplane is rotated for takeoff; or
    (4) The warning is manually deactivated by the pilot.
    (c) The means used to activate the system must function properly 
throughout the ranges of takeoff weights, altitudes, and temperatures 
for which certification is requested.

[Amdt. 25-42, 43 FR 2323, Jan. 16, 1978]

                              Landing Gear



Sec. 25.721  General.

    (a) The landing gear system must be designed so that when it fails 
due to overloads during takeoff and landing, the failure mode is not 
likely to cause spillage of enough fuel to constitute a fire hazard. The 
overloads must be assumed to act in the upward and aft directions in 
combination with side loads acting inboard and outboard. In the absence 
of a more rational analysis, the side loads must be assumed to be up to 
20 percent of the vertical load or 20 percent of the drag load, 
whichever is greater.
    (b) The airplane must be designed to avoid any rupture leading to 
the spillage of enough fuel to constitute a fire hazard as a result of a 
wheels-up landing on a paved runway, under the following minor crash 
landing conditions:
    (1) Impact at 5 feet-per-second vertical velocity, with the airplane 
under control, at Maximum Design Landing Weight--
    (i) With the landing gear fully retracted; and
    (ii) With any one or more landing gear legs not extended.
    (2) Sliding on the ground, with--
    (i) The landing gear fully retracted and with up to a 20[deg] yaw 
angle; and
    (ii) Any one or more landing gear legs not extended and with 0[deg] 
yaw angle.
    (c) For configurations where the engine nacelle is likely to come 
into contact with the ground, the engine pylon or engine mounting must 
be designed so that when it fails due to overloads (assuming the 
overloads to act predominantly in the upward direction and separately, 
predominantly in the aft direction), the failure mode is not likely to 
cause the spillage of enough fuel to constitute a fire hazard.

[Amdt. No. 25-139, 79 FR 59430, Oct. 2, 2014]



Sec. 25.723  Shock absorption tests.

    (a) The analytical representation of the landing gear dynamic 
characteristics that is used in determining the

[[Page 442]]

landing loads must be validated by energy absorption tests. A range of 
tests must be conducted to ensure that the analytical representation is 
valid for the design conditions specified in Sec. 25.473.
    (1) The configurations subjected to energy absorption tests at limit 
design conditions must include at least the design landing weight or the 
design takeoff weight, whichever produces the greater value of landing 
impact energy.
    (2) The test attitude of the landing gear unit and the application 
of appropriate drag loads during the test must simulate the airplane 
landing conditions in a manner consistent with the development of 
rational or conservative limit loads.
    (b) The landing gear may not fail in a test, demonstrating its 
reserve energy absorption capacity, simulating a descent velocity of 12 
f.p.s. at design landing weight, assuming airplane lift not greater than 
airplane weight acting during the landing impact.
    (c) In lieu of the tests prescribed in this section, changes in 
previously approved design weights and minor changes in design may be 
substantiated by analyses based on previous tests conducted on the same 
basic landing gear system that has similar energy absorption 
characteristics.

[Doc. No. 1999-5835, 66 FR 27394, May 16, 2001]



Sec. Sec. 25.725-25.727  [Reserved]



Sec. 25.729  Retracting mechanism.

    (a) General. For airplanes with retractable landing gear, the 
following apply:
    (1) The landing gear retracting mechanism, wheel well doors, and 
supporting structure, must be designed for--
    (i) The loads occurring in the flight conditions when the gear is in 
the retracted position,
    (ii) The combination of friction loads, inertia loads, brake torque 
loads, air loads, and gyroscopic loads resulting from the wheels 
rotating at a peripheral speed equal to 1.23VSR (with the 
wing-flaps in take-off position at design take-off weight), occurring 
during retraction and extension at any airspeed up to 1.5 
VSR1 (with the wing-flaps in the approach position at design 
landing weight), and
    (iii) Any load factor up to those specified in Sec. 25.345(a) for 
the wing-flaps extended condition.
    (2) Unless there are other means to decelerate the airplane in 
flight at this speed, the landing gear, the retracting mechanism, and 
the airplane structure (including wheel well doors) must be designed to 
withstand the flight loads occurring with the landing gear in the 
extended position at any speed up to 0.67 VC.
    (3) Landing gear doors, their operating mechanism, and their 
supporting structures must be designed for the yawing maneuvers 
prescribed for the airplane in addition to the conditions of airspeed 
and load factor prescribed in paragraphs (a)(1) and (2) of this section.
    (b) Landing gear lock. There must be positive means to keep the 
landing gear extended in flight and on the ground. There must be 
positive means to keep the landing gear and doors in the correct 
retracted position in flight, unless it can be shown that lowering of 
the landing gear or doors, or flight with the landing gear or doors 
extended, at any speed, is not hazardous.
    (c) Emergency operation. There must be an emergency means for 
extending the landing gear in the event of--
    (1) Any reasonably probable failure in the normal retraction system; 
or
    (2) The failure of any single source of hydraulic, electric, or 
equivalent energy supply.
    (d) Operation test. The proper functioning of the retracting 
mechanism must be shown by operation tests.
    (e) Position indicator and warning device. If a retractable landing 
gear is used, there must be a landing gear position indicator easily 
visible to the pilot or to the appropriate crew members (as well as 
necessary devices to actuate the indicator) to indicate without 
ambiguity that the retractable units and their associated doors are 
secured in the extended (or retracted) position. The means must be 
designed as follows:
    (1) If switches are used, they must be located and coupled to the 
landing gear mechanical systems in a manner that prevents an erroneous 
indication of

[[Page 443]]

``down and locked'' if the landing gear is not in a fully extended 
position, or of ``up and locked'' if the landing gear is not in the 
fully retracted position. The switches may be located where they are 
operated by the actual landing gear locking latch or device.
    (2) The flightcrew must be given an aural warning that functions 
continuously, or is periodically repeated, if a landing is attempted 
when the landing gear is not locked down.
    (3) The warning must be given in sufficient time to allow the 
landing gear to be locked down or a go-around to be made.
    (4) There must not be a manual shut-off means readily available to 
the flightcrew for the warning required by paragraph (e)(2) of this 
section such that it could be operated instinctively, inadvertently, or 
by habitual reflexive action.
    (5) The system used to generate the aural warning must be designed 
to minimize false or inappropriate alerts.
    (6) Failures of systems used to inhibit the landing gear aural 
warning, that would prevent the warning system from operating, must be 
improbable.
    (7) A flightcrew alert must be provided whenever the landing gear 
position is not consistent with the landing gear selector lever 
position.
    (f) Protection of equipment on landing gear and in wheel wells. 
Equipment that is essential to the safe operation of the airplane and 
that is located on the landing gear and in wheel wells must be protected 
from the damaging effects of--
    (1) A bursting tire;
    (2) A loose tire tread, unless it is shown that a loose tire tread 
cannot cause damage.
    (3) Possible wheel brake temperatures.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5676, Apr. 8, 1970; Amdt. 25-42, 43 FR 2323, Jan. 16, 1978; Amdt. 
25-72, 55 FR 29777, July 20, 1990; Amdt. 25-75, 56 FR 63762, Dec. 5, 
1991; Amdt. 25-136, 77 FR 1617, Jan. 11, 2012]



Sec. 25.731  Wheels.

    (a) Each main and nose wheel must be approved.
    (b) The maximum static load rating of each wheel may not be less 
than the corresponding static ground reaction with--
    (1) Design maximum weight; and
    (2) Critical center of gravity.
    (c) The maximum limit load rating of each wheel must equal or exceed 
the maximum radial limit load determined under the applicable ground 
load requirements of this part.
    (d) Overpressure burst prevention. Means must be provided in each 
wheel to prevent wheel failure and tire burst that may result from 
excessive pressurization of the wheel and tire assembly.
    (e) Braked wheels. Each braked wheel must meet the applicable 
requirements of Sec. 25.735.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29777, July 20, 1990; Amdt. 25-107, 67 FR 20420, Apr. 24, 2002]



Sec. 25.733  Tires.

    (a) When a landing gear axle is fitted with a single wheel and tire 
assembly, the wheel must be fitted with a suitable tire of proper fit 
with a speed rating approved by the Administrator that is not exceeded 
under critical conditions and with a load rating approved by the 
Administrator that is not exceeded under--
    (1) The loads on the main wheel tire, corresponding to the most 
critical combination of airplane weight (up to maximum weight) and 
center of gravity position, and
    (2) The loads corresponding to the ground reactions in paragraph (b) 
of this section, on the nose wheel tire, except as provided in 
paragraphs (b)(2) and (b)(3) of this section.
    (b) The applicable ground reactions for nose wheel tires are as 
follows:
    (1) The static ground reaction for the tire corresponding to the 
most critical combination of airplane weight (up to maximum ramp weight) 
and center of gravity position with a force of 1.0g acting downward at 
the center of gravity. This load may not exceed the load rating of the 
tire.
    (2) The ground reaction of the tire corresponding to the most 
critical combination of airplane weight (up to maximum landing weight) 
and center of gravity position combined with forces of 1.0g downward and 
0.31g forward acting at the center of gravity.

[[Page 444]]

The reactions in this case must be distributed to the nose and main 
wheels by the principles of statics with a drag reaction equal to 0.31 
times the vertical load at each wheel with brakes capable of producing 
this ground reaction. This nose tire load may not exceed 1.5 times the 
load rating of the tire.
    (3) The ground reaction of the tire corresponding to the most 
critical combination of airplane weight (up to maximum ramp weight) and 
center of gravity position combined with forces of 1.0g downward and 
0.20g forward acting at the center of gravity. The reactions in this 
case must be distributed to the nose and main wheels by the principles 
of statics with a drag reaction equal to 0.20 times the vertical load at 
each wheel with brakes capable of producing this ground reaction. This 
nose tire load may not exceed 1.5 times the load rating of the tire.
    (c) When a landing gear axle is fitted with more than one wheel and 
tire assembly, such as dual or dual-tandem, each wheel must be fitted 
with a suitable tire of proper fit with a speed rating approved by the 
Administrator that is not exceeded under critical conditions, and with a 
load rating approved by the Administrator that is not exceeded by--
    (1) The loads on each main wheel tire, corresponding to the most 
critical combination of airplane weight (up to maximum weight) and 
center of gravity position, when multiplied by a factor of 1.07; and
    (2) Loads specified in paragraphs (a)(2), (b)(1), (b)(2), and (b)(3) 
of this section on each nose wheel tire.
    (d) Each tire installed on a retractable landing gear system must, 
at the maximum size of the tire type expected in service, have a 
clearance to surrounding structure and systems that is adequate to 
prevent unintended contact between the tire and any part of the 
structure or systems.
    (e) For an airplane with a maximum certificated takeoff weight of 
more than 75,000 pounds, tires mounted on braked wheels must be inflated 
with dry nitrogen or other gases shown to be inert so that the gas 
mixture in the tire does not contain oxygen in excess of 5 percent by 
volume, unless it can be shown that the tire liner material will not 
produce a volatile gas when heated or that means are provided to prevent 
tire temperatures from reaching unsafe levels.

[Amdt. 25-48, 44 FR 68752, Nov. 29, 1979; Amdt. 25-72, 55 FR 29777, July 
20, 1990, as amended by Amdt. 25-78, 58 FR 11781, Feb. 26, 1993]



Sec. 25.735  Brakes and braking systems.

    (a) Approval. Each assembly consisting of a wheel(s) and brake(s) 
must be approved.
    (b) Brake system capability. The brake system, associated systems 
and components must be designed and constructed so that:
    (1) If any electrical, pneumatic, hydraulic, or mechanical 
connecting or transmitting element fails, or if any single source of 
hydraulic or other brake operating energy supply is lost, it is possible 
to bring the airplane to rest with a braked roll stopping distance of 
not more than two times that obtained in determining the landing 
distance as prescribed in Sec. 25.125.
    (2) Fluid lost from a brake hydraulic system following a failure in, 
or in the vicinity of, the brakes is insufficient to cause or support a 
hazardous fire on the ground or in flight.
    (c) Brake controls. The brake controls must be designed and 
constructed so that:
    (1) Excessive control force is not required for their operation.
    (2) If an automatic braking system is installed, means are provided 
to:
    (i) Arm and disarm the system, and
    (ii) Allow the pilot(s) to override the system by use of manual 
braking.
    (d) Parking brake. The airplane must have a parking brake control 
that, when selected on, will, without further attention, prevent the 
airplane from rolling on a dry and level paved runway when the most 
adverse combination of maximum thrust on one engine and up to maximum 
ground idle thrust on any, or all, other engine(s) is applied. The 
control must be suitably located or be adequately protected to prevent 
inadvertent operation. There must be indication in the cockpit when the 
parking brake is not fully released.
    (e) Antiskid system. If an antiskid system is installed:

[[Page 445]]

    (1) It must operate satisfactorily over the range of expected runway 
conditions, without external adjustment.
    (2) It must, at all times, have priority over the automatic braking 
system, if installed.
    (f) Kinetic energy capacity--(1) Design landing stop. The design 
landing stop is an operational landing stop at maximum landing weight. 
The design landing stop brake kinetic energy absorption requirement of 
each wheel, brake, and tire assembly must be determined. It must be 
substantiated by dynamometer testing that the wheel, brake and tire 
assembly is capable of absorbing not less than this level of kinetic 
energy throughout the defined wear range of the brake. The energy 
absorption rate derived from the airplane manufacturer's braking 
requirements must be achieved. The mean deceleration must not be less 
than 10 fps \2\.
    (2) Maximum kinetic energy accelerate-stop. The maximum kinetic 
energy accelerate-stop is a rejected takeoff for the most critical 
combination of airplane takeoff weight and speed. The accelerate-stop 
brake kinetic energy absorption requirement of each wheel, brake, and 
tire assembly must be determined. It must be substantiated by 
dynamometer testing that the wheel, brake, and tire assembly is capable 
of absorbing not less than this level of kinetic energy throughout the 
defined wear range of the brake. The energy absorption rate derived from 
the airplane manufacturer's braking requirements must be achieved. The 
mean deceleration must not be less than 6 fps\2\.
    (3) Most severe landing stop. The most severe landing stop is a stop 
at the most critical combination of airplane landing weight and speed. 
The most severe landing stop brake kinetic energy absorption requirement 
of each wheel, brake, and tire assembly must be determined. It must be 
substantiated by dynamometer testing that, at the declared fully worn 
limit(s) of the brake heat sink, the wheel, brake and tire assembly is 
capable of absorbing not less than this level of kinetic energy. The 
most severe landing stop need not be considered for extremely improbable 
failure conditions or if the maximum kinetic energy accelerate-stop 
energy is more severe.
    (g) Brake condition after high kinetic energy dynamometer stop(s). 
Following the high kinetic energy stop demonstration(s) required by 
paragraph (f) of this section, with the parking brake promptly and fully 
applied for at least 3 minutes, it must be demonstrated that for at 
least 5 minutes from application of the parking brake, no condition 
occurs (or has occurred during the stop), including fire associated with 
the tire or wheel and brake assembly, that could prejudice the safe and 
complete evacuation of the airplane.
    (h) Stored energy systems. An indication to the flightcrew of the 
usable stored energy must be provided if a stored energy system is used 
to show compliance with paragraph (b)(1) of this section. The available 
stored energy must be sufficient for:
    (1) At least 6 full applications of the brakes when an antiskid 
system is not operating; and
    (2) Bringing the airplane to a complete stop when an antiskid system 
is operating, under all runway surface conditions for which the airplane 
is certificated.
    (i) Brake wear indicators. Means must be provided for each brake 
assembly to indicate when the heat sink is worn to the permissible 
limit. The means must be reliable and readily visible.
    (j) Overtemperature burst prevention. Means must be provided in each 
braked wheel to prevent a wheel failure, a tire burst, or both, that may 
result from elevated brake temperatures. Additionally, all wheels must 
meet the requirements of Sec. 25.731(d).
    (k) Compatibility. Compatibility of the wheel and brake assemblies 
with the airplane and its systems must be substantiated.

[Doc. No. FAA-1999-6063, 67 FR 20420, Apr. 24, 2002, as amended by Amdt. 
25-108, 67 FR 70827, Nov. 26, 2002; 68 FR 1955, Jan. 15, 2003]



Sec. 25.737  Skis.

    Each ski must be approved. The maximum limit load rating of each ski 
must equal or exceed the maximum limit load determined under the 
applicable ground load requirements of this part.

[[Page 446]]

                            Floats and Hulls



Sec. 25.751  Main float buoyancy.

    Each main float must have--
    (a) A buoyancy of 80 percent in excess of that required to support 
the maximum weight of the seaplane or amphibian in fresh water; and
    (b) Not less than five watertight compartments approximately equal 
in volume.



Sec. 25.753  Main float design.

    Each main float must be approved and must meet the requirements of 
Sec. 25.521.



Sec. 25.755  Hulls.

    (a) Each hull must have enough watertight compartments so that, with 
any two adjacent compartments flooded, the buoyancy of the hull and 
auxiliary floats (and wheel tires, if used) provides a margin of 
positive stability great enough to minimize the probability of capsizing 
in rough, fresh water.
    (b) Bulkheads with watertight doors may be used for communication 
between compartments.

                   Personnel and Cargo Accommodations



Sec. 25.771  Pilot compartment.

    (a) Each pilot compartment and its equipment must allow the minimum 
flight crew (established under Sec. 25.1523) to perform their duties 
without unreasonable concentration or fatigue.
    (b) The primary controls listed in Sec. 25.779(a), excluding cables 
and control rods, must be located with respect to the propellers so that 
no member of the minimum flight crew (established under Sec. 25.1523), 
or part of the controls, lies in the region between the plane of 
rotation of any inboard propeller and the surface generated by a line 
passing through the center of the propeller hub making an angle of five 
degrees forward or aft of the plane of rotation of the propeller.
    (c) If provision is made for a second pilot, the airplane must be 
controllable with equal safety from either pilot seat.
    (d) The pilot compartment must be constructed so that, when flying 
in rain or snow, it will not leak in a manner that will distract the 
crew or harm the structure.
    (e) Vibration and noise characteristics of cockpit equipment may not 
interfere with safe operation of the airplane.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-4, 30 
FR 6113, Apr. 30, 1965]



Sec. 25.772  Pilot compartment doors.

    For an airplane that has a lockable door installed between the pilot 
compartment and the passenger compartment:
    (a) For airplanes with a maximum passenger seating configuration of 
more than 20 seats, the emergency exit configuration must be designed so 
that neither crewmembers nor passengers require use of the flightdeck 
door in order to reach the emergency exits provided for them; and
    (b) Means must be provided to enable flight crewmembers to directly 
enter the passenger compartment from the pilot compartment if the 
cockpit door becomes jammed.
    (c) There must be an emergency means to enable a flight attendant to 
enter the pilot compartment in the event that the flightcrew becomes 
incapacitated.

[Doc. No. 24344, 55 FR 29777, July 20, 1990, as amended by Amdt. 25-106, 
67 FR 2127, Jan. 15, 2002]



Sec. 25.773  Pilot compartment view.

    (a) Nonprecipitation conditions. For nonprecipitation conditions, 
the following apply:
    (1) Each pilot compartment must be arranged to give the pilots a 
sufficiently extensive, clear, and undistorted view, to enable them to 
safely perform any maneuvers within the operating limitations of the 
airplane, including taxiing takeoff, approach, and landing.
    (2) Each pilot compartment must be free of glare and reflection that 
could interfere with the normal duties of the minimum flight crew 
(established under Sec. 25.1523). This must be shown in day and night 
flight tests under nonprecipitation conditions.

[[Page 447]]

    (b) Precipitation conditions. For precipitation conditions, the 
following apply:
    (1) The airplane must have a means to maintain a clear portion of 
the windshield, during precipitation conditions, sufficient for both 
pilots to have a sufficiently extensive view along the flight path in 
normal flight attitudes of the airplane. This means must be designed to 
function, without continuous attention on the part of the crew, in--
    (i) Heavy rain at speeds up to 1.5 VSR1 with lift and 
drag devices retracted; and
    (ii) The icing conditions specified in Sec. 25.1419 if 
certification for flight in icing conditions is requested.
    (2) No single failure of the systems used to provide the view 
required by paragraph (b)(1) of this section must cause the loss of that 
view by both pilots in the specified precipitation conditions.
    (3) The first pilot must have a window that--
    (i) Is openable under the conditions prescribed in paragraph (b)(1) 
of this section when the cabin is not pressurized;
    (ii) Provides the view specified in paragraph (b)(1) of this 
section; and
    (iii) Provides sufficient protection from the elements against 
impairment of the pilot's vision.
    (4) The openable window specified in paragraph (b)(3) of this 
section need not be provided if it is shown that an area of the 
transparent surface will remain clear sufficient for at least one pilot 
to land the airplane safely in the event of--
    (i) Any system failure or combination of failures which is not 
extremely improbable, in accordance with Sec. 25.1309, under the 
precipitation conditions specified in paragraph (b)(1) of this section.
    (ii) An encounter with severe hail, birds, or insects.
    (c) Internal windshield and window fogging. The airplane must have a 
means to prevent fogging of the internal portions of the windshield and 
window panels over an area which would provide the visibility specified 
in paragraph (a) of this section under all internal and external ambient 
conditions, including precipitation conditions, in which the airplane is 
intended to be operated.
    (d) Fixed markers or other guides must be installed at each pilot 
station to enable the pilots to position themselves in their seats for 
an optimum combination of outside visibility and instrument scan. If 
lighted markers or guides are used they must comply with the 
requirements specified in Sec. 25.1381.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5676, Apr. 8, 1970; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 
25-72, 55 FR 29778, July 20, 1990; Amdt. 25-108, 67 FR 70827, Nov. 26, 
2002; Amdt. 25-121, 72 FR 44669, Aug. 8, 2007; Amdt. 25-136, 77 FR 1618, 
Jan. 11, 2012]

    Effective Date Note: At 79 FR 65526, Nov. 4, 2014, Sec. 25.773 was 
amended by revising paragraph (b)(1)(ii), effective Jan. 5, 2015. For 
the convenience of the user, the revised text is set forth as follows:



Sec. 25.773  Pilot compartment view.

                                * * * * *

    (b) * * *
    (1) * * *
    (ii) The icing conditions specified in Appendix C of this part and 
the following icing conditions specified in Appendix O of this part, if 
certification for flight in icing conditions is sought:
    (A) For airplanes certificated in accordance with Sec. 
25.1420(a)(1), the icing conditions that the airplane is certified to 
safely exit following detection.
    (B) For airplanes certificated in accordance with Sec. 
25.1420(a)(2), the icing conditions that the airplane is certified to 
safely operate in and the icing conditions that the airplane is 
certified to safely exit following detection.
    (C) For airplanes certificated in accordance with Sec. 
25.1420(a)(3) and for airplanes not subject to Sec. 25.1420, all icing 
conditions.

                                * * * * *



Sec. 25.775  Windshields and windows.

    (a) Internal panes must be made of nonsplintering material.
    (b) Windshield panes directly in front of the pilots in the normal 
conduct of their duties, and the supporting structures for these panes, 
must withstand, without penetration, the impact of a four-pound bird 
when the velocity of the airplane (relative to the bird along the 
airplane's flight path) is equal to

[[Page 448]]

the value of VC, at sea level, selected under Sec. 25.335(a).
    (c) Unless it can be shown by analysis or tests that the probability 
of occurrence of a critical windshield fragmentation condition is of a 
low order, the airplane must have a means to minimize the danger to the 
pilots from flying windshield fragments due to bird impact. This must be 
shown for each transparent pane in the cockpit that--
    (1) Appears in the front view of the airplane;
    (2) Is inclined 15 degrees or more to the longitudinal axis of the 
airplane; and
    (3) Has any part of the pane located where its fragmentation will 
constitute a hazard to the pilots.
    (d) The design of windshields and windows in pressurized airplanes 
must be based on factors peculiar to high altitude operation, including 
the effects of continuous and cyclic pressurization loadings, the 
inherent characteristics of the material used, and the effects of 
temperatures and temperature differentials. The windshield and window 
panels must be capable of withstanding the maximum cabin pressure 
differential loads combined with critical aerodynamic pressure and 
temperature effects after any single failure in the installation or 
associated systems. It may be assumed that, after a single failure that 
is obvious to the flight crew (established under Sec. 25.1523), the 
cabin pressure differential is reduced from the maximum, in accordance 
with appropriate operating limitations, to allow continued safe flight 
of the airplane with a cabin pressure altitude of not more than 15,000 
feet.
    (e) The windshield panels in front of the pilots must be arranged so 
that, assuming the loss of vision through any one panel, one or more 
panels remain available for use by a pilot seated at a pilot station to 
permit continued safe flight and landing.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5676, Apr. 8, 1970; Amdt. 25-38, 41 FR 55466, Dec. 20, 1976]



Sec. 25.777  Cockpit controls.

    (a) Each cockpit control must be located to provide convenient 
operation and to prevent confusion and inadvertent operation.
    (b) The direction of movement of cockpit controls must meet the 
requirements of Sec. 25.779. Wherever practicable, the sense of motion 
involved in the operation of other controls must correspond to the sense 
of the effect of the operation upon the airplane or upon the part 
operated. Controls of a variable nature using a rotary motion must move 
clockwise from the off position, through an increasing range, to the 
full on position.
    (c) The controls must be located and arranged, with respect to the 
pilots' seats, so that there is full and unrestricted movement of each 
control without interference from the cockpit structure or the clothing 
of the minimum flight crew (established under Sec. 25.1523) when any 
member of this flight crew, from 5[foot]2 to 
6[foot]3 in height, is seated with the seat belt and shoulder 
harness (if provided) fastened.
    (d) Identical powerplant controls for each engine must be located to 
prevent confusion as to the engines they control.
    (e) Wing flap controls and other auxiliary lift device controls must 
be located on top of the pedestal, aft of the throttles, centrally or to 
the right of the pedestal centerline, and not less than 10 inches aft of 
the landing gear control.
    (f) The landing gear control must be located forward of the 
throttles and must be operable by each pilot when seated with seat belt 
and shoulder harness (if provided) fastened.
    (g) Control knobs must be shaped in accordance with Sec. 25.781. In 
addition, the knobs must be of the same color, and this color must 
contrast with the color of control knobs for other purposes and the 
surrounding cockpit.
    (h) If a flight engineer is required as part of the minimum flight 
crew (established under Sec. 25.1523), the airplane must have a flight 
engineer station located and arranged so that the flight crewmembers can 
perform their functions efficiently and without interfering with each 
other.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 
43 FR 50596, Oct. 30, 1978]

[[Page 449]]



Sec. 25.779  Motion and effect of cockpit controls.

    Cockpit controls must be designed so that they operate in accordance 
with the following movement and actuation:
    (a) Aerodynamic controls:
    (1) Primary.

------------------------------------------------------------------------
                 Controls                        Motion and effect
------------------------------------------------------------------------
Aileron..................................  Right (clockwise) for right
                                            wing down.
Elevator.................................  Rearward for nose up.
Rudder...................................  Right pedal forward for nose
                                            right.
------------------------------------------------------------------------

    (2) Secondary.

------------------------------------------------------------------------
                 Controls                        Motion and effect
------------------------------------------------------------------------
Flaps (or auxiliary lift devices)........  Forward for flaps up;
                                            rearward for flaps down.
Trim tabs (or equivalent)................  Rotate to produce similar
                                            rotation of the airplane
                                            about an axis parallel to
                                            the axis of the control.
------------------------------------------------------------------------

    (b) Powerplant and auxiliary controls:
    (1) Powerplant.

------------------------------------------------------------------------
                 Controls                        Motion and effect
------------------------------------------------------------------------
Power or thrust..........................  Forward to increase forward
                                            thrust and rearward to
                                            increase rearward thrust.
Propellers...............................  Forward to increase rpm.
Mixture..................................  Forward or upward for rich.
Carburetor air heat......................  Forward or upward for cold.
Supercharger.............................  Forward or upward for low
                                            blower. For
                                            turbosuperchargers, forward,
                                            upward, or clockwise, to
                                            increase pressure.
------------------------------------------------------------------------

    (2) Auxiliary.

------------------------------------------------------------------------
                 Controls                        Motion and effect
------------------------------------------------------------------------
Landing gear.............................  Down to extend.
------------------------------------------------------------------------


[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29778, July 20, 1990]



Sec. 25.781  Cockpit control knob shape.

    Cockpit control knobs must conform to the general shapes (but not 
necessarily the exact sizes or specific proportions) in the following 
figure:

[[Page 450]]

[GRAPHIC] [TIFF OMITTED] TC28SE91.048


[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29779, July 20, 1990]



Sec. 25.783  Fuselage doors.

    (a) General. This section applies to fuselage doors, which includes 
all doors, hatches, openable windows, access panels, covers, etc., on 
the exterior of the fuselage that do not require the use of tools to 
open or close. This also applies to each door or hatch through a 
pressure bulkhead, including any bulkhead that is specifically designed 
to function as a secondary bulkhead under the prescribed failure 
conditions of part 25. These doors must meet the requirements of this 
section, taking into account both pressurized and unpressurized flight, 
and must be designed as follows:
    (1) Each door must have means to safeguard against opening in flight 
as a result of mechanical failure, or failure of any single structural 
element.

[[Page 451]]

    (2) Each door that could be a hazard if it unlatches must be 
designed so that unlatching during pressurized and unpressurized flight 
from the fully closed, latched, and locked condition is extremely 
improbable. This must be shown by safety analysis.
    (3) Each element of each door operating system must be designed or, 
where impracticable, distinctively and permanently marked, to minimize 
the probability of incorrect assembly and adjustment that could result 
in a malfunction.
    (4) All sources of power that could initiate unlocking or unlatching 
of any door must be automatically isolated from the latching and locking 
systems prior to flight and it must not be possible to restore power to 
the door during flight.
    (5) Each removable bolt, screw, nut, pin, or other removable 
fastener must meet the locking requirements of Sec. 25.607.
    (6) Certain doors, as specified by Sec. 25.807(h), must also meet 
the applicable requirements of Sec. Sec. 25.809 through 25.812 for 
emergency exits.
    (b) Opening by persons. There must be a means to safeguard each door 
against opening during flight due to inadvertent action by persons. In 
addition, design precautions must be taken to minimize the possibility 
for a person to open a door intentionally during flight. If these 
precautions include the use of auxiliary devices, those devices and 
their controlling systems must be designed so that--
    (1) No single failure will prevent more than one exit from being 
opened; and
    (2) Failures that would prevent opening of the exit after landing 
are improbable.
    (c) Pressurization prevention means. There must be a provision to 
prevent pressurization of the airplane to an unsafe level if any door 
subject to pressurization is not fully closed, latched, and locked.
    (1) The provision must be designed to function after any single 
failure, or after any combination of failures not shown to be extremely 
improbable.
    (2) Doors that meet the conditions described in paragraph (h) of 
this section are not required to have a dedicated pressurization 
prevention means if, from every possible position of the door, it will 
remain open to the extent that it prevents pressurization or safely 
close and latch as pressurization takes place. This must also be shown 
with any single failure and malfunction, except that--
    (i) With failures or malfunctions in the latching mechanism, it need 
not latch after closing; and
    (ii) With jamming as a result of mechanical failure or blocking 
debris, the door need not close and latch if it can be shown that the 
pressurization loads on the jammed door or mechanism would not result in 
an unsafe condition.
    (d) Latching and locking. The latching and locking mechanisms must 
be designed as follows:
    (1) There must be a provision to latch each door.
    (2) The latches and their operating mechanism must be designed so 
that, under all airplane flight and ground loading conditions, with the 
door latched, there is no force or torque tending to unlatch the 
latches. In addition, the latching system must include a means to secure 
the latches in the latched position. This means must be independent of 
the locking system.
    (3) Each door subject to pressurization, and for which the initial 
opening movement is not inward, must--
    (i) Have an individual lock for each latch;
    (ii) Have the lock located as close as practicable to the latch; and
    (iii) Be designed so that, during pressurized flight, no single 
failure in the locking system would prevent the locks from restraining 
the latches necessary to secure the door.
    (4) Each door for which the initial opening movement is inward, and 
unlatching of the door could result in a hazard, must have a locking 
means to prevent the latches from becoming disengaged. The locking means 
must ensure sufficient latching to prevent opening of the door even with 
a single failure of the latching mechanism.
    (5) It must not be possible to position the lock in the locked 
position if the latch and the latching mechanism are not in the latched 
position.

[[Page 452]]

    (6) It must not be possible to unlatch the latches with the locks in 
the locked position. Locks must be designed to withstand the limit loads 
resulting from--
    (i) The maximum operator effort when the latches are operated 
manually;
    (ii) The powered latch actuators, if installed; and
    (iii) The relative motion between the latch and the structural 
counterpart.
    (7) Each door for which unlatching would not result in a hazard is 
not required to have a locking mechanism meeting the requirements of 
paragraphs (d)(3) through (d)(6) of this section.
    (e) Warning, caution, and advisory indications. Doors must be 
provided with the following indications:
    (1) There must be a positive means to indicate at each door 
operator's station that all required operations to close, latch, and 
lock the door(s) have been completed.
    (2) There must be a positive means clearly visible from each 
operator station for any door that could be a hazard if unlatched to 
indicate if the door is not fully closed, latched, and locked.
    (3) There must be a visual means on the flight deck to signal the 
pilots if any door is not fully closed, latched, and locked. The means 
must be designed such that any failure or combination of failures that 
would result in an erroneous closed, latched, and locked indication is 
improbable for--
    (i) Each door that is subject to pressurization and for which the 
initial opening movement is not inward; or
    (ii) Each door that could be a hazard if unlatched.
    (4) There must be an aural warning to the pilots prior to or during 
the initial portion of takeoff roll if any door is not fully closed, 
latched, and locked, and its opening would prevent a safe takeoff and 
return to landing.
    (f) Visual inspection provision. Each door for which unlatching of 
the door could be a hazard must have a provision for direct visual 
inspection to determine, without ambiguity, if the door is fully closed, 
latched, and locked. The provision must be permanent and discernible 
under operational lighting conditions, or by means of a flashlight or 
equivalent light source.
    (g) Certain maintenance doors, removable emergency exits, and access 
panels. Some doors not normally opened except for maintenance purposes 
or emergency evacuation and some access panels need not comply with 
certain paragraphs of this section as follows:
    (1) Access panels that are not subject to cabin pressurization and 
would not be a hazard if open during flight need not comply with 
paragraphs (a) through (f) of this section, but must have a means to 
prevent inadvertent opening during flight.
    (2) Inward-opening removable emergency exits that are not normally 
removed, except for maintenance purposes or emergency evacuation, and 
flight deck-openable windows need not comply with paragraphs (c) and (f) 
of this section.
    (3) Maintenance doors that meet the conditions of paragraph (h) of 
this section, and for which a placard is provided limiting use to 
maintenance access, need not comply with paragraphs (c) and (f) of this 
section.
    (h) Doors that are not a hazard. For the purposes of this section, a 
door is considered not to be a hazard in the unlatched condition during 
flight, provided it can be shown to meet all of the following 
conditions:
    (1) Doors in pressurized compartments would remain in the fully 
closed position if not restrained by the latches when subject to a 
pressure greater than \1/2\ psi. Opening by persons, either 
inadvertently or intentionally, need not be considered in making this 
determination.
    (2) The door would remain inside the airplane or remain attached to 
the airplane if it opens either in pressurized or unpressurized portions 
of the flight. This determination must include the consideration of 
inadvertent and intentional opening by persons during either pressurized 
or unpressurized portions of the flight.
    (3) The disengagement of the latches during flight would not allow 
depressurization of the cabin to an unsafe level. This safety assessment 
must include the physiological effects on the occupants.

[[Page 453]]

    (4) The open door during flight would not create aerodynamic 
interference that could preclude safe flight and landing.
    (5) The airplane would meet the structural design requirements with 
the door open. This assessment must include the aeroelastic stability 
requirements of Sec. 25.629, as well as the strength requirements of 
subpart C of this part.
    (6) The unlatching or opening of the door must not preclude safe 
flight and landing as a result of interaction with other systems or 
structures.

[Doc. No. 2003-14193, 69 FR 24501, May 3, 2004]



Sec. 25.785  Seats, berths, safety belts, and harnesses.

    (a) A seat (or berth for a nonambulant person) must be provided for 
each occupant who has reached his or her second birthday.
    (b) Each seat, berth, safety belt, harness, and adjacent part of the 
airplane at each station designated as occupiable during takeoff and 
landing must be designed so that a person making proper use of these 
facilities will not suffer serious injury in an emergency landing as a 
result of the inertia forces specified in Sec. Sec. 25.561 and 25.562.
    (c) Each seat or berth must be approved.
    (d) Each occupant of a seat that makes more than an 18-degree angle 
with the vertical plane containing the airplane centerline must be 
protected from head injury by a safety belt and an energy absorbing rest 
that will support the arms, shoulders, head, and spine, or by a safety 
belt and shoulder harness that will prevent the head from contacting any 
injurious object. Each occupant of any other seat must be protected from 
head injury by a safety belt and, as appropriate to the type, location, 
and angle of facing of each seat, by one or more of the following:
    (1) A shoulder harness that will prevent the head from contacting 
any injurious object.
    (2) The elimination of any injurious object within striking radius 
of the head.
    (3) An energy absorbing rest that will support the arms, shoulders, 
head, and spine.
    (e) Each berth must be designed so that the forward part has a 
padded end board, canvas diaphragm, or equivalent means, that can 
withstand the static load reaction of the occupant when subjected to the 
forward inertia force specified in Sec. 25.561. Berths must be free 
from corners and protuberances likely to cause injury to a person 
occupying the berth during emergency conditions.
    (f) Each seat or berth, and its supporting structure, and each 
safety belt or harness and its anchorage must be designed for an 
occupant weight of 170 pounds, considering the maximum load factors, 
inertia forces, and reactions among the occupant, seat, safety belt, and 
harness for each relevant flight and ground load condition (including 
the emergency landing conditions prescribed in Sec. 25.561). In 
addition--
    (1) The structural analysis and testing of the seats, berths, and 
their supporting structures may be determined by assuming that the 
critical load in the forward, sideward, downward, upward, and rearward 
directions (as determined from the prescribed flight, ground, and 
emergency landing conditions) acts separately or using selected 
combinations of loads if the required strength in each specified 
direction is substantiated. The forward load factor need not be applied 
to safety belts for berths.
    (2) Each pilot seat must be designed for the reactions resulting 
from the application of the pilot forces prescribed in Sec. 25.395.
    (3) The inertia forces specified in Sec. 25.561 must be multiplied 
by a factor of 1.33 (instead of the fitting factor prescribed in Sec. 
25.625) in determining the strength of the attachment of each seat to 
the structure and each belt or harness to the seat or structure.
    (g) Each seat at a flight deck station must have a restraint system 
consisting of a combined safety belt and shoulder harness with a single-
point release that permits the flight deck occupant, when seated with 
the restraint system fastened, to perform all of the occupant's 
necessary flight deck functions. There must be a means to secure each 
combined restraint system when not in use to prevent interference with 
the operation of the airplane and with rapid egress in an emergency.

[[Page 454]]

    (h) Each seat located in the passenger compartment and designated 
for use during takeoff and landing by a flight attendant required by the 
operating rules of this chapter must be:
    (1) Near a required floor level emergency exit, except that another 
location is acceptable if the emergency egress of passengers would be 
enhanced with that location. A flight attendant seat must be located 
adjacent to each Type A or B emergency exit. Other flight attendant 
seats must be evenly distributed among the required floor- level 
emergency exits to the extent feasible.
    (2) To the extent possible, without compromising proximity to a 
required floor level emergency exit, located to provide a direct view of 
the cabin area for which the flight attendant is responsible.
    (3) Positioned so that the seat will not interfere with the use of a 
passageway or exit when the seat is not in use.
    (4) Located to minimize the probability that occupants would suffer 
injury by being struck by items dislodged from service areas, stowage 
compartments, or service equipment.
    (5) Either forward or rearward facing with an energy absorbing rest 
that is designed to support the arms, shoulders, head, and spine.
    (6) Equipped with a restraint system consisting of a combined safety 
belt and shoulder harness unit with a single point release. There must 
be means to secure each restraint system when not in use to prevent 
interference with rapid egress in an emergency.
    (i) Each safety belt must be equipped with a metal to metal latching 
device.
    (j) If the seat backs do not provide a firm handhold, there must be 
a handgrip or rail along each aisle to enable persons to steady 
themselves while using the aisles in moderately rough air.
    (k) Each projecting object that would injure persons seated or 
moving about the airplane in normal flight must be padded.
    (l) Each forward observer's seat required by the operating rules 
must be shown to be suitable for use in conducting the necessary enroute 
inspection.

[Amdt. 25-72, 55 FR 29780, July 20, 1990, as amended by Amdt. 25-88, 61 
FR 57956, Nov. 8, 1996]



Sec. 25.787  Stowage compartments.

    (a) Each compartment for the stowage of cargo, baggage, carry-on 
articles, and equipment (such as life rafts), and any other stowage 
compartment, must be designed for its placarded maximum weight of 
contents and for the critical load distribution at the appropriate 
maximum load factors corresponding to the specified flight and ground 
load conditions, and to those emergency landing conditions of Sec. 
25.561(b)(3) for which the breaking loose of the contents of such 
compartments in the specified direction could--
    (1) Cause direct injury to occupants;
    (2) Penetrate fuel tanks or lines or cause fire or explosion hazard 
by damage to adjacent systems; or
    (3) Nullify any of the escape facilities provided for use after an 
emergency landing.

If the airplane has a passenger-seating configuration, excluding pilot 
seats, of 10 seats or more, each stowage compartment in the passenger 
cabin, except for under seat and overhead compartments for passenger 
convenience, must be completely enclosed.
    (b) There must be a means to prevent the contents in the 
compartments from becoming a hazard by shifting, under the loads 
specified in paragraph (a) of this section. For stowage compartments in 
the passenger and crew cabin, if the means used is a latched door, the 
design must take into consideration the wear and deterioration expected 
in service.
    (c) If cargo compartment lamps are installed, each lamp must be 
installed so as to prevent contact between lamp bulb and cargo.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-32, 
37 FR 3969, Feb. 24, 1972; Amdt. 25-38, 41 FR 55466, Dec. 20, 1976; 
Amdt. 25-51, 45 FR 7755, Feb. 4, 1980; Amdt. No. 25-139, 79 FR 59430, 
Oct. 2, 2014]

[[Page 455]]



Sec. 25.789  Retention of items of mass in passenger and crew
compartments and galleys.

    (a) Means must be provided to prevent each item of mass (that is 
part of the airplane type design) in a passenger or crew compartment or 
galley from becoming a hazard by shifting under the appropriate maximum 
load factors corresponding to the specified flight and ground load 
conditions, and to the emergency landing conditions of Sec. 25.561(b).
    (b) Each interphone restraint system must be designed so that when 
subjected to the load factors specified in Sec. 25.561(b)(3), the 
interphone will remain in its stowed position.

[Amdt. 25-32, 37 FR 3969, Feb. 24, 1972, as amended by Amdt. 25-46, 43 
FR 50596, Oct. 30, 1978]



Sec. 25.791  Passenger information signs and placards.

    (a) If smoking is to be prohibited, there must be at least one 
placard so stating that is legible to each person seated in the cabin. 
If smoking is to be allowed, and if the crew compartment is separated 
from the passenger compartment, there must be at least one sign 
notifying when smoking is prohibited. Signs which notify when smoking is 
prohibited must be operable by a member of the flightcrew and, when 
illuminated, must be legible under all probable conditions of cabin 
illumination to each person seated in the cabin.
    (b) Signs that notify when seat belts should be fastened and that 
are installed to comply with the operating rules of this chapter must be 
operable by a member of the flightcrew and, when illuminated, must be 
legible under all probable conditions of cabin illumination to each 
person seated in the cabin.
    (c) A placard must be located on or adjacent to the door of each 
receptacle used for the disposal of flammable waste materials to 
indicate that use of the receptacle for disposal of cigarettes, etc., is 
prohibited.
    (d) Lavatories must have ``No Smoking'' or ``No Smoking in 
Lavatory'' placards conspicuously located on or adjacent to each side of 
the entry door.
    (e) Symbols that clearly express the intent of the sign or placard 
may be used in lieu of letters.

[Amdt. 25-72, 55 FR 29780, July 20, 1990]



Sec. 25.793  Floor surfaces.

    The floor surface of all areas which are likely to become wet in 
service must have slip resistant properties.

[Amdt. 25-51, 45 FR 7755, Feb. 4, 1980]



Sec. 25.795  Security considerations.

    (a) Protection of flightcrew compartment. If a flightdeck door is 
required by operating rules:
    (1) The bulkhead, door, and any other accessible boundary separating 
the flightcrew compartment from occupied areas must be designed to 
resist forcible intrusion by unauthorized persons and be capable of 
withstanding impacts of 300 joules (221.3 foot pounds).
    (2) The bulkhead, door, and any other accessible boundary separating 
the flightcrew compartment from occupied areas must be designed to 
resist a constant 250 pound (1,113 Newtons) tensile load on accessible 
handholds, including the doorknob or handle.
    (3) The bulkhead, door, and any other boundary separating the 
flightcrew compartment from any occupied areas must be designed to 
resist penetration by small arms fire and fragmentation devices to a 
level equivalent to level IIIa of the National Institute of Justice 
(NIJ) Standard 0101.04.
    (b) Airplanes with a maximum certificated passenger seating capacity 
of more than 60 persons or a maximum certificated takeoff gross weight 
of over 100,000 pounds (45,359 Kilograms) must be designed to limit the 
effects of an explosive or incendiary device as follows:
    (1) Flightdeck smoke protection. Means must be provided to limit 
entry of smoke, fumes, and noxious gases into the flightdeck.
    (2) Passenger cabin smoke protection. Means must be provided to 
prevent passenger incapacitation in the cabin resulting from smoke, 
fumes, and noxious gases as represented by the initial combined 
volumetric concentrations of 0.59% carbon monoxide and 1.23% carbon 
dioxide.

[[Page 456]]

    (3) Cargo compartment fire suppression. An extinguishing agent must 
be capable of suppressing a fire. All cargo-compartment fire suppression 
systems must be designed to withstand the following effects, including 
support structure displacements or adjacent materials displacing against 
the distribution system:
    (i) Impact or damage from a 0.5-inch diameter aluminum sphere 
traveling at 430 feet per second (131.1 meters per second);
    (ii) A 15-pound per square-inch (103.4 kPa) pressure load if the 
projected surface area of the component is greater than 4 square feet. 
Any single dimension greater than 4 feet (1.22 meters) may be assumed to 
be 4 feet (1.22 meters) in length; and
    (iii) A 6-inch (0.152 meters) displacement, except where limited by 
the fuselage contour, from a single point force applied anywhere along 
the distribution system where relative movement between the system and 
its attachment can occur.
    (iv) Paragraphs (b)(3)(i) through (iii) of this section do not apply 
to components that are redundant and separated in accordance with 
paragraph (c)(2) of this section or are installed remotely from the 
cargo compartment.
    (c) An airplane with a maximum certificated passenger seating 
capacity of more than 60 persons or a maximum certificated takeoff gross 
weight of over 100,000 pounds (45,359 Kilograms) must comply with the 
following:
    (1) Least risk bomb location. An airplane must be designed with a 
designated location where a bomb or other explosive device could be 
placed to best protect flight-critical structures and systems from 
damage in the case of detonation.
    (2) Survivability of systems. (i) Except where impracticable, 
redundant airplane systems necessary for continued safe flight and 
landing must be physically separated, at a minimum, by an amount equal 
to a sphere of diameter
[GRAPHIC] [TIFF OMITTED] TR28OC08.003


(where H0 is defined under Sec. 25.365(e)(2) of this part 
and D need not exceed 5.05 feet (1.54 meters)). The sphere is applied 
everywhere within the fuselage--limited by the forward bulkhead and the 
aft bulkhead of the passenger cabin and cargo compartment beyond which 
only one-half the sphere is applied.
    (ii) Where compliance with paragraph (c)(2)(i) of this section is 
impracticable, other design precautions must be taken to maximize the 
survivability of those systems.
    (3) Interior design to facilitate searches. Design features must be 
incorporated that will deter concealment or promote discovery of 
weapons, explosives, or other objects from a simple inspection in the 
following areas of the airplane cabin:
    (i) Areas above the overhead bins must be designed to prevent 
objects from being hidden from view in a simple search from the aisle. 
Designs that prevent concealment of objects with volumes 20 cubic inches 
and greater satisfy this requirement.
    (ii) Toilets must be designed to prevent the passage of solid 
objects greater than 2.0 inches in diameter.
    (iii) Life preservers or their storage locations must be designed so 
that tampering is evident.
    (d) Each chemical oxygen generator or its installation must be 
designed to be secure from deliberate manipulation by one of the 
following:
    (1) By providing effective resistance to tampering,
    (2) By providing an effective combination of resistance to tampering 
and active tamper-evident features,
    (3) By installation in a location or manner whereby any attempt to 
access the generator would be immediately obvious, or
    (4) By a combination of approaches specified in paragraphs (d)(1), 
(d)(2) and (d)(3) of this section that the Administrator finds provides 
a secure installation.
    (e) Exceptions. Airplanes used solely to transport cargo only need 
to meet the requirements of paragraphs (b)(1), (b)(3), and (c)(2) of 
this section.
    (f) Material Incorporated by Reference. You must use National 
Institute of Justice (NIJ) Standard 0101.04, Ballistic Resistance of 
Personal Body Armor, June 2001, Revision A, to establish ballistic 
resistance as required by paragraph (a)(3) of this section.

[[Page 457]]

    (1) The Director of the Federal Register approved the incorporation 
by reference of this document under 5 U.S.C. 552(a) and 1 CFR part 51.
    (2) You may review copies of NIJ Standard 0101.04 at the:
    (i) FAA Transport Airplane Directorate, 1601 Lind Avenue, SW., 
Renton, Washington 98055;
    (ii) National Institute of Justice (NIJ), http://www.ojp.usdoj.gov/
nij, telephone (202) 307-2942; or
    (iii) National Archives and Records Administration (NARA). For 
information on the availability of this material at NARA go to http://
www.archives.gov/federal--register/code--of--federal--regulations/ibr--
locations.html or call (202) 741-6030.
    (3) You may obtain copies of NIJ Standard 0101.04 from the National 
Criminal Justice Reference Service, P.O. Box 6000, Rockville, MD 20849-
6000, telephone (800) 851-3420.

[Amdt. Nos. 25-127; 121-341, 73 FR 63879, Oct. 28, 2008, as amended at 
74 FR 22819, May 15, 2009; Amdt. 25-138, 79 FR 13519, Mar. 11, 2014]

                          Emergency Provisions



Sec. 25.801  Ditching.

    (a) If certification with ditching provisions is requested, the 
airplane must meet the requirements of this section and Sec. Sec. 
25.807(e), 25.1411, and 25.1415(a).
    (b) Each practicable design measure, compatible with the general 
characteristics of the airplane, must be taken to minimize the 
probability that in an emergency landing on water, the behavior of the 
airplane would cause immediate injury to the occupants or would make it 
impossible for them to escape.
    (c) The probable behavior of the airplane in a water landing must be 
investigated by model tests or by comparison with airplanes of similar 
configuration for which the ditching characteristics are known. Scoops, 
flaps, projections, and any other factor likely to affect the 
hydrodynamic characteristics of the airplane, must be considered.
    (d) It must be shown that, under reasonably probable water 
conditions, the flotation time and trim of the airplane will allow the 
occupants to leave the airplane and enter the liferafts required by 
Sec. 25.1415. If compliance with this provision is shown by buoyancy 
and trim computations, appropriate allowances must be made for probable 
structural damage and leakage. If the airplane has fuel tanks (with fuel 
jettisoning provisions) that can reasonably be expected to withstand a 
ditching without leakage, the jettisonable volume of fuel may be 
considered as buoyancy volume.
    (e) Unless the effects of the collapse of external doors and windows 
are accounted for in the investigation of the probable behavior of the 
airplane in a water landing (as prescribed in paragraphs (c) and (d) of 
this section), the external doors and windows must be designed to 
withstand the probable maximum local pressures.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29781, July 20, 1990]



Sec. 25.803  Emergency evacuation.

    (a) Each crew and passenger area must have emergency means to allow 
rapid evacuation in crash landings, with the landing gear extended as 
well as with the landing gear retracted, considering the possibility of 
the airplane being on fire.
    (b) [Reserved]
    (c) For airplanes having a seating capacity of more than 44 
passengers, it must be shown that the maximum seating capacity, 
including the number of crewmembers required by the operating rules for 
which certification is requested, can be evacuated from the airplane to 
the ground under simulated emergency conditions within 90 seconds. 
Compliance with this requirement must be shown by actual demonstration 
using the test criteria outlined in appendix J of this part unless the 
Administrator finds that a combination of analysis and testing will 
provide data equivalent to that which would be obtained by actual 
demonstration.
    (d)-(e) [Reserved]

[Doc. No. 24344, 55 FR 29781, July 20, 1990]



Sec. 25.807  Emergency exits.

    (a) Type. For the purpose of this part, the types of exits are 
defined as follows:

[[Page 458]]

    (1) Type I. This type is a floor-level exit with a rectangular 
opening of not less than 24 inches wide by 48 inches high, with corner 
radii not greater than eight inches.
    (2) Type II. This type is a rectangular opening of not less than 20 
inches wide by 44 inches high, with corner radii not greater than seven 
inches. Type II exits must be floor-level exits unless located over the 
wing, in which case they must not have a step-up inside the airplane of 
more than 10 inches nor a step-down outside the airplane of more than 17 
inches.
    (3) Type III. This type is a rectangular opening of not less than 20 
inches wide by 36 inches high with corner radii not greater than seven 
inches, and with a step-up inside the airplane of not more than 20 
inches. If the exit is located over the wing, the step-down outside the 
airplane may not exceed 27 inches.
    (4) Type IV. This type is a rectangular opening of not less than 19 
inches wide by 26 inches high, with corner radii not greater than 6.3 
inches, located over the wing, with a step-up inside the airplane of not 
more than 29 inches and a step-down outside the airplane of not more 
than 36 inches.
    (5) Ventral. This type is an exit from the passenger compartment 
through the pressure shell and the bottom fuselage skin. The dimensions 
and physical configuration of this type of exit must allow at least the 
same rate of egress as a Type I exit with the airplane in the normal 
ground attitude, with landing gear extended.
    (6) Tailcone. This type is an aft exit from the passenger 
compartment through the pressure shell and through an openable cone of 
the fuselage aft of the pressure shell. The means of opening the 
tailcone must be simple and obvious and must employ a single operation.
    (7) Type A. This type is a floor-level exit with a rectangular 
opening of not less than 42 inches wide by 72 inches high, with corner 
radii not greater than seven inches.
    (8) Type B. This type is a floor-level exit with a rectangular 
opening of not less than 32 inches wide by 72 inches high, with corner 
radii not greater than six inches.
    (9) Type C. This type is a floor-level exit with a rectangular 
opening of not less than 30 inches wide by 48 inches high, with corner 
radii not greater than 10 inches.
    (b) Step down distance. Step down distance, as used in this section, 
means the actual distance between the bottom of the required opening and 
a usable foot hold, extending out from the fuselage, that is large 
enough to be effective without searching by sight or feel.
    (c) Over-sized exits. Openings larger than those specified in this 
section, whether or not of rectangular shape, may be used if the 
specified rectangular opening can be inscribed within the opening and 
the base of the inscribed rectangular opening meets the specified step-
up and step-down heights.
    (d) Asymmetry. Exits of an exit pair need not be diametrically 
opposite each other nor of the same size; however, the number of 
passenger seats permitted under paragraph (g) of this section is based 
on the smaller of the two exits.
    (e) Uniformity. Exits must be distributed as uniformly as practical, 
taking into account passenger seat distribution.
    (f) Location. (1) Each required passenger emergency exit must be 
accessible to the passengers and located where it will afford the most 
effective means of passenger evacuation.
    (2) If only one floor-level exit per side is prescribed, and the 
airplane does not have a tailcone or ventral emergency exit, the floor-
level exits must be in the rearward part of the passenger compartment 
unless another location affords a more effective means of passenger 
evacuation.
    (3) If more than one floor-level exit per side is prescribed, and 
the airplane does not have a combination cargo and passenger 
configuration, at least one floor-level exit must be located in each 
side near each end of the cabin.
    (4) For an airplane that is required to have more than one passenger 
emergency exit for each side of the fuselage, no passenger emergency 
exit shall be more than 60 feet from any adjacent passenger emergency 
exit on the same side of the same deck of the fuselage,

[[Page 459]]

as measured parallel to the airplane's longitudinal axis between the 
nearest exit edges.
    (g) Type and number required. The maximum number of passenger seats 
permitted depends on the type and number of exits installed in each side 
of the fuselage. Except as further restricted in paragraphs (g)(1) 
through (g)(9) of this section, the maximum number of passenger seats 
permitted for each exit of a specific type installed in each side of the 
fuselage is as follows:

Type A                                                           110
Type B                                                            75
Type C                                                            55
Type I                                                            45
Type II                                                           40
Type III                                                          35
Type IV                                                            9
 

    (1) For a passenger seating configuration of 1 to 9 seats, there 
must be at least one Type IV or larger overwing exit in each side of the 
fuselage or, if overwing exits are not provided, at least one exit in 
each side that meets the minimum dimensions of a Type III exit.
    (2) For a passenger seating configuration of more than 9 seats, each 
exit must be a Type III or larger exit.
    (3) For a passenger seating configuration of 10 to 19 seats, there 
must be at least one Type III or larger exit in each side of the 
fuselage.
    (4) For a passenger seating configuration of 20 to 40 seats, there 
must be at least two exits, one of which must be a Type II or larger 
exit, in each side of the fuselage.
    (5) For a passenger seating configuration of 41 to 110 seats, there 
must be at least two exits, one of which must be a Type I or larger 
exit, in each side of the fuselage.
    (6) For a passenger seating configuration of more than 110 seats, 
the emergency exits in each side of the fuselage must include at least 
two Type I or larger exits.
    (7) The combined maximum number of passenger seats permitted for all 
Type III exits is 70, and the combined maximum number of passenger seats 
permitted for two Type III exits in each side of the fuselage that are 
separated by fewer than three passenger seat rows is 65.
    (8) If a Type A, Type B, or Type C exit is installed, there must be 
at least two Type C or larger exits in each side of the fuselage.
    (9) If a passenger ventral or tailcone exit is installed and that 
exit provides at least the same rate of egress as a Type III exit with 
the airplane in the most adverse exit opening condition that would 
result from the collapse of one or more legs of the landing gear, an 
increase in the passenger seating configuration is permitted as follows:
    (i) For a ventral exit, 12 additional passenger seats.
    (ii) For a tailcone exit incorporating a floor level opening of not 
less than 20 inches wide by 60 inches high, with corner radii not 
greater than seven inches, in the pressure shell and incorporating an 
approved assist means in accordance with Sec. 25.810(a), 25 additional 
passenger seats.
    (iii) For a tailcone exit incorporating an opening in the pressure 
shell which is at least equivalent to a Type III emergency exit with 
respect to dimensions, step-up and step-down distance, and with the top 
of the opening not less than 56 inches from the passenger compartment 
floor, 15 additional passenger seats.
    (h) Other exits. The following exits also must meet the applicable 
emergency exit requirements of Sec. Sec. 25.809 through 25.812, and 
must be readily accessible:
    (1) Each emergency exit in the passenger compartment in excess of 
the minimum number of required emergency exits.
    (2) Any other floor-level door or exit that is accessible from the 
passenger compartment and is as large or larger than a Type II exit, but 
less than 46 inches wide.
    (3) Any other ventral or tail cone passenger exit.
    (i) Ditching emergency exits for passengers. Whether or not ditching 
certification is requested, ditching emergency exits must be provided in 
accordance with the following requirements, unless the emergency exits 
required by paragraph (g) of this section already meet them:
    (1) For airplanes that have a passenger seating configuration of 
nine or fewer seats, excluding pilot seats, one

[[Page 460]]

exit above the waterline in each side of the airplane, meeting at least 
the dimensions of a Type IV exit.
    (2) For airplanes that have a passenger seating configuration of 10 
of more seats, excluding pilot seats, one exit above the waterline in a 
side of the airplane, meeting at least the dimensions of a Type III exit 
for each unit (or part of a unit) of 35 passenger seats, but no less 
than two such exits in the passenger cabin, with one on each side of the 
airplane. The passenger seat/ exit ratio may be increased through the 
use of larger exits, or other means, provided it is shown that the 
evacuation capability during ditching has been improved accordingly.
    (3) If it is impractical to locate side exits above the waterline, 
the side exits must be replaced by an equal number of readily accessible 
overhead hatches of not less than the dimensions of a Type III exit, 
except that for airplanes with a passenger configuration of 35 or fewer 
seats, excluding pilot seats, the two required Type III side exits need 
be replaced by only one overhead hatch.
    (j) Flightcrew emergency exits. For airplanes in which the proximity 
of passenger emergency exits to the flightcrew area does not offer a 
convenient and readily accessible means of evacuation of the flightcrew, 
and for all airplanes having a passenger seating capacity greater than 
20, flightcrew exits shall be located in the flightcrew area. Such exits 
shall be of sufficient size and so located as to permit rapid evacuation 
by the crew. One exit shall be provided on each side of the airplane; 
or, alternatively, a top hatch shall be provided. Each exit must 
encompass an unobstructed rectangular opening of at least 19 by 20 
inches unless satisfactory exit utility can be demonstrated by a typical 
crewmember.

[Amdt. 25-72, 55 FR 29781, July 20, 1990, as amended by Amdt. 25-88, 61 
FR 57956, Nov. 8, 1996; 62 FR 1817, Jan. 13, 1997; Amdt. 25-94, 63 FR 
8848, Feb. 23, 1998; 63 FR 12862, Mar. 16, 1998; Amdt. 25-114, 69 FR 
24502, May 3, 2004]



Sec. 25.809  Emergency exit arrangement.

    (a) Each emergency exit, including each flightcrew emergency exit, 
must be a moveable door or hatch in the external walls of the fuselage, 
allowing an unobstructed opening to the outside. In addition, each 
emergency exit must have means to permit viewing of the conditions 
outside the exit when the exit is closed. The viewing means may be on or 
adjacent to the exit provided no obstructions exist between the exit and 
the viewing means. Means must also be provided to permit viewing of the 
likely areas of evacuee ground contact. The likely areas of evacuee 
ground contact must be viewable during all lighting conditions with the 
landing gear extended as well as in all conditions of landing gear 
collapse.
    (b) Each emergency exit must be openable from the inside and the 
outside except that sliding window emergency exits in the flight crew 
area need not be openable from the outside if other approved exits are 
convenient and readily accessible to the flight crew area. Each 
emergency exit must be capable of being opened, when there is no 
fuselage deformation--
    (1) With the airplane in the normal ground attitude and in each of 
the attitudes corresponding to collapse of one or more legs of the 
landing gear; and
    (2) Within 10 seconds measured from the time when the opening means 
is actuated to the time when the exit is fully opened.
    (3) Even though persons may be crowded against the door on the 
inside of the airplane.
    (c) The means of opening emergency exits must be simple and obvious; 
may not require exceptional effort; and must be arranged and marked so 
that it can be readily located and operated, even in darkness. Internal 
exit-opening means involving sequence operations (such as operation of 
two handles or latches, or the release of safety catches) may be used 
for flightcrew emergency exits if it can be reasonably established that 
these means are simple and obvious to crewmembers trained in their use.
    (d) If a single power-boost or single power-operated system is the 
primary system for operating more than one exit in an emergency, each 
exit must be capable of meeting the requirements of paragraph (b) of 
this section in the event of failure of the primary system.

[[Page 461]]

Manual operation of the exit (after failure of the primary system) is 
acceptable.
    (e) Each emergency exit must be shown by tests, or by a combination 
of analysis and tests, to meet the requirements of paragraphs (b) and 
(c) of this section.
    (f) Each door must be located where persons using them will not be 
endangered by the propellers when appropriate operating procedures are 
used.
    (g) There must be provisions to minimize the probability of jamming 
of the emergency exits resulting from fuselage deformation in a minor 
crash landing.
    (h) When required by the operating rules for any large passenger-
carrying turbojet-powered airplane, each ventral exit and tailcone exit 
must be--
    (1) Designed and constructed so that it cannot be opened during 
flight; and
    (2) Marked with a placard readable from a distance of 30 inches and 
installed at a conspicuous location near the means of opening the exit, 
stating that the exit has been designed and constructed so that it 
cannot be opened during flight.
    (i) Each emergency exit must have a means to retain the exit in the 
open position, once the exit is opened in an emergency. The means must 
not require separate action to engage when the exit is opened, and must 
require positive action to disengage.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-15, 
32 FR 13264, Sept. 20, 1967; Amdt. 25-32, 37 FR 3970, Feb. 24, 1972; 
Amdt. 25-34, 37 FR 25355, Nov. 30, 1972; Amdt. 25-46, 43 FR 50597, Oct. 
30, 1978; Amdt. 25-47, 44 FR 61325, Oct. 25, 1979; Amdt. 25-72, 55 FR 
29782, July 20, 1990; Amdt. 25-114, 69 FR 24502, May 3, 2004; Amdt. 25-
116, 69 FR 62788, Oct. 27, 2004]



Sec. 25.810  Emergency egress assist means and escape routes.

    (a) Each non over-wing Type A, Type B or Type C exit, and any other 
non over-wing landplane emergency exit more than 6 feet from the ground 
with the airplane on the ground and the landing gear extended, must have 
an approved means to assist the occupants in descending to the ground.
    (1) The assisting means for each passenger emergency exit must be a 
self-supporting slide or equivalent; and, in the case of Type A or Type 
B exits, it must be capable of carrying simultaneously two parallel 
lines of evacuees. In addition, the assisting means must be designed to 
meet the following requirements--
    (i) It must be automatically deployed and deployment must begin 
during the interval between the time the exit opening means is actuated 
from inside the airplane and the time the exit is fully opened. However, 
each passenger emergency exit which is also a passenger entrance door or 
a service door must be provided with means to prevent deployment of the 
assisting means when it is opened from either the inside or the outside 
under nonemergency conditions for normal use.
    (ii) Except for assisting means installed at Type C exits, it must 
be automatically erected within 6 seconds after deployment is begun. 
Assisting means installed at Type C exits must be automatically erected 
within 10 seconds from the time the opening means of the exit is 
actuated.
    (iii) It must be of such length after full deployment that the lower 
end is self-supporting on the ground and provides safe evacuation of 
occupants to the ground after collapse of one or more legs of the 
landing gear.
    (iv) It must have the capability, in 25-knot winds directed from the 
most critical angle, to deploy and, with the assistance of only one 
person, to remain usable after full deployment to evacuate occupants 
safely to the ground.
    (v) For each system installation (mockup or airplane installed), 
five consecutive deployment and inflation tests must be conducted (per 
exit) without failure, and at least three tests of each such five-test 
series must be conducted using a single representative sample of the 
device. The sample devices must be deployed and inflated by the system's 
primary means after being subjected to the inertia forces specified in 
Sec. 25.561(b). If any part of the system fails or does not function 
properly during the required tests, the cause of the failure or 
malfunction must be corrected by positive means and after that, the full 
series of five consecutive deployment and inflation

[[Page 462]]

tests must be conducted without failure.
    (2) The assisting means for flightcrew emergency exits may be a rope 
or any other means demonstrated to be suitable for the purpose. If the 
assisting means is a rope, or an approved device equivalent to a rope, 
it must be--
    (i) Attached to the fuselage structure at or above the top of the 
emergency exit opening, or, for a device at a pilot's emergency exit 
window, at another approved location if the stowed device, or its 
attachment, would reduce the pilot's view in flight;
    (ii) Able (with its attachment) to withstand a 400-pound static 
load.
    (b) Assist means from the cabin to the wing are required for each 
type A or Type B exit located above the wing and having a stepdown 
unless the exit without an assist-means can be shown to have a rate of 
passenger egress at least equal to that of the same type of non over-
wing exit. If an assist means is required, it must be automatically 
deployed and automatically erected concurrent with the opening of the 
exit. In the case of assist means installed at Type C exits, it must be 
self-supporting within 10 seconds from the time the opening means of the 
exits is actuated. For all other exit types, it must be self-supporting 
6 seconds after deployment is begun.
    (c) An escape route must be established from each overwing emergency 
exit, and (except for flap surfaces suitable as slides) covered with a 
slip resistant surface. Except where a means for channeling the flow of 
evacuees is provided--
    (1) The escape route from each Type A or Type B passenger emergency 
exit, or any common escape route from two Type III passenger emergency 
exits, must be at least 42 inches wide; that from any other passenger 
emergency exit must be at least 24 inches wide; and
    (2) The escape route surface must have a reflectance of at least 80 
percent, and must be defined by markings with a surface-to-marking 
contrast ratio of at least 5:1.
    (d) Means must be provided to assist evacuees to reach the ground 
for all Type C exits located over the wing and, if the place on the 
airplane structure at which the escape route required in paragraph (c) 
of this section terminates is more than 6 feet from the ground with the 
airplane on the ground and the landing gear extended, for all other exit 
types.
    (1) If the escape route is over the flap, the height of the terminal 
edge must be measured with the flap in the takeoff or landing position, 
whichever is higher from the ground.
    (2) The assisting means must be usable and self-supporting with one 
or more landing gear legs collapsed and under a 25-knot wind directed 
from the most critical angle.
    (3) The assisting means provided for each escape route leading from 
a Type A or B emergency exit must be capable of carrying simultaneously 
two parallel lines of evacuees; and, the assisting means leading from 
any other exit type must be capable of carrying as many parallel lines 
of evacuees as there are required escape routes.
    (4) The assisting means provided for each escape route leading from 
a Type C exit must be automatically erected within 10 seconds from the 
time the opening means of the exit is actuated, and that provided for 
the escape route leading from any other exit type must be automatically 
erected within 10 seconds after actuation of the erection system.
    (e) If an integral stair is installed in a passenger entry door that 
is qualified as a passenger emergency exit, the stair must be designed 
so that, under the following conditions, the effectiveness of passenger 
emergency egress will not be impaired:
    (1) The door, integral stair, and operating mechanism have been 
subjected to the inertia forces specified in Sec. 25.561(b)(3), acting 
separately relative to the surrounding structure.
    (2) The airplane is in the normal ground attitude and in each of the 
attitudes corresponding to collapse of one or more legs of the landing 
gear.

[Amdt. 25-72, 55 FR 29782, July 20, 1990, as amended by Amdt. 25-88, 61 
FR 57958, Nov. 8, 1996; 62 FR 1817, Jan. 13, 1997; Amdt. 25-114, 69 FR 
24502, May 3, 2004]



Sec. 25.811  Emergency exit marking.

    (a) Each passenger emergency exit, its means of access, and its 
means of

[[Page 463]]

opening must be conspicuously marked.
    (b) The identity and location of each passenger emergency exit must 
be recognizable from a distance equal to the width of the cabin.
    (c) Means must be provided to assist the occupants in locating the 
exits in conditions of dense smoke.
    (d) The location of each passenger emergency exit must be indicated 
by a sign visible to occupants approaching along the main passenger 
aisle (or aisles). There must be--
    (1) A passenger emergency exit locator sign above the aisle (or 
aisles) near each passenger emergency exit, or at another overhead 
location if it is more practical because of low headroom, except that 
one sign may serve more than one exit if each exit can be seen readily 
from the sign;
    (2) A passenger emergency exit marking sign next to each passenger 
emergency exit, except that one sign may serve two such exits if they 
both can be seen readily from the sign; and
    (3) A sign on each bulkhead or divider that prevents fore and aft 
vision along the passenger cabin to indicate emergency exits beyond and 
obscured by the bulkhead or divider, except that if this is not possible 
the sign may be placed at another appropriate location.
    (e) The location of the operating handle and instructions for 
opening exits from the inside of the airplane must be shown in the 
following manner:
    (1) Each passenger emergency exit must have, on or near the exit, a 
marking that is readable from a distance of 30 inches.
    (2) Each Type A, Type B, Type C or Type I passenger emergency exit 
operating handle must--
    (i) Be self-illuminated with an initial brightness of at least 160 
microlamberts; or
    (ii) Be conspicuously located and well illuminated by the emergency 
lighting even in conditions of occupant crowding at the exit.
    (3) [Reserved]
    (4) Each Type A, Type B, Type C, Type I, or Type II passenger 
emergency exit with a locking mechanism released by rotary motion of the 
handle must be marked--
    (i) With a red arrow, with a shaft at least three-fourths of an inch 
wide and a head twice the width of the shaft, extending along at least 
70 degrees of arc at a radius approximately equal to three-fourths of 
the handle length.
    (ii) So that the centerline of the exit handle is within 1 inch of the projected point of the arrow when the 
handle has reached full travel and has released the locking mechanism, 
and
    (iii) With the word ``open'' in red letters 1 inch high, placed 
horizontally near the head of the arrow.
    (f) Each emergency exit that is required to be openable from the 
outside, and its means of opening, must be marked on the outside of the 
airplane. In addition, the following apply:
    (1) The outside marking for each passenger emergency exit in the 
side of the fuselage must include a 2-inch colored band outlining the 
exit.
    (2) Each outside marking including the band, must have color 
contrast to be readily distinguishable from the surrounding fuselage 
surface. The contrast must be such that if the reflectance of the darker 
color is 15 percent or less, the reflectance of the lighter color must 
be at least 45 percent. ``Reflectance'' is the ratio of the luminous 
flux reflected by a body to the luminous flux it receives. When the 
reflectance of the darker color is greater than 15 percent, at least a 
30-percent difference between its reflectance and the reflectance of the 
lighter color must be provided.
    (3) In the case of exists other than those in the side of the 
fuselage, such as ventral or tailcone exists, the external means of 
opening, including instructions if applicable, must be conspicuously 
marked in red, or bright chrome yellow if the background color is such 
that red is inconspicuous. When the opening means is located on only one 
side of the fuselage, a conspicuous marking to that effect must be 
provided on the other side.
    (g) Each sign required by paragraph (d) of this section may use the 
word

[[Page 464]]

``exit'' in its legend in place of the term ``emergency exit''.

[Amdt. 25-15, 32 FR 13264, Sept. 20, 1967, as amended by Amdt. 25-32, 37 
FR 3970, Feb. 24, 1972; Amdt. 25-46, 43 FR 50597, Oct. 30, 1978; 43 FR 
52495, Nov. 13, 1978; Amdt. 25-79, 58 FR 45229, Aug. 26, 1993; Amdt. 25-
88, 61 FR 57958, Nov. 8, 1996]



Sec. 25.812  Emergency lighting.

    (a) An emergency lighting system, independent of the main lighting 
system, must be installed. However, the sources of general cabin 
illumination may be common to both the emergency and the main lighting 
systems if the power supply to the emergency lighting system is 
independent of the power supply to the main lighting system. The 
emergency lighting system must include:
    (1) Illuminated emergency exit marking and locating signs, sources 
of general cabin illumination, interior lighting in emergency exit 
areas, and floor proximity escape path marking.
    (2) Exterior emergency lighting.
    (b) Emergency exit signs--
    (1) For airplanes that have a passenger seating configuration, 
excluding pilot seats, of 10 seats or more must meet the following 
requirements:
    (i) Each passenger emergency exit locator sign required by Sec. 
25.811(d)(1) and each passenger emergency exit marking sign required by 
Sec. 25.811(d)(2) must have red letters at least 1\1/2\ inches high on 
an illuminated white background, and must have an area of at least 21 
square inches excluding the letters. The lighted background-to-letter 
contrast must be at least 10:1. The letter height to stroke-width ratio 
may not be more than 7:1 nor less than 6:1. These signs must be 
internally electrically illuminated with a background brightness of at 
least 25 foot-lamberts and a high-to-low background contrast no greater 
than 3:1.
    (ii) Each passenger emergency exit sign required by Sec. 
25.811(d)(3) must have red letters at least 1\1/2\ inches high on a 
white background having an area of at least 21 square inches excluding 
the letters. These signs must be internally electrically illuminated or 
self-illuminated by other than electrical means and must have an initial 
brightness of at least 400 microlamberts. The colors may be reversed in 
the case of a sign that is self-illuminated by other than electrical 
means.
    (2) For airplanes that have a passenger seating configuration, 
excluding pilot seats, of nine seats or less, that are required by Sec. 
25.811(d)(1), (2), and (3) must have red letters at least 1 inch high on 
a white background at least 2 inches high. These signs may be internally 
electrically illuminated, or self-illuminated by other than electrical 
means, with an initial brightness of at least 160 microlamberts. The 
colors may be reversed in the case of a sign that is self-illuminated by 
other than electrical means.
    (c) General illumination in the passenger cabin must be provided so 
that when measured along the centerline of main passenger aisle(s), and 
cross aisle(s) between main aisles, at seat arm-rest height and at 40-
inch intervals, the average illumination is not less than 0.05 foot-
candle and the illumination at each 40-inch interval is not less than 
0.01 foot-candle. A main passenger aisle(s) is considered to extend 
along the fuselage from the most forward passenger emergency exit or 
cabin occupant seat, whichever is farther forward, to the most rearward 
passenger emergency exit or cabin occupant seat, whichever is farther 
aft.
    (d) The floor of the passageway leading to each floor-level 
passenger emergency exit, between the main aisles and the exit openings, 
must be provided with illumination that is not less than 0.02 foot-
candle measured along a line that is within 6 inches of and parallel to 
the floor and is centered on the passenger evacuation path.
    (e) Floor proximity emergency escape path marking must provide 
emergency evacuation guidance for passengers when all sources of 
illumination more than 4 feet above the cabin aisle floor are totally 
obscured. In the dark of the night, the floor proximity emergency escape 
path marking must enable each passenger to--
    (1) After leaving the passenger seat, visually identify the 
emergency escape path along the cabin aisle floor to the first exits or 
pair of exits forward and aft of the seat; and

[[Page 465]]

    (2) Readily identify each exit from the emergency escape path by 
reference only to markings and visual features not more than 4 feet 
above the cabin floor.
    (f) Except for subsystems provided in accordance with paragraph (h) 
of this section that serve no more than one assist means, are 
independent of the airplane's main emergency lighting system, and are 
automatically activated when the assist means is erected, the emergency 
lighting system must be designed as follows.
    (1) The lights must be operable manually from the flight crew 
station and from a point in the passenger compartment that is readily 
accessible to a normal flight attendant seat.
    (2) There must be a flight crew warning light which illuminates when 
power is on in the airplane and the emergency lighting control device is 
not armed.
    (3) The cockpit control device must have an ``on,'' ``off,'' and 
``armed'' position so that when armed in the cockpit or turned on at 
either the cockpit or flight attendant station the lights will either 
light or remain lighted upon interruption (except an interruption caused 
by a transverse vertical separation of the fuselage during crash 
landing) of the airplane's normal electric power. There must be a means 
to safeguard against inadvertent operation of the control device from 
the ``armed'' or ``on'' positions.
    (g) Exterior emergency lighting must be provided as follows:
    (1) At each overwing emergency exit the illumination must be--
    (i) Not less than 0.03 foot-candle (measured normal to the direction 
of the incident light) on a 2-square-foot area where an evacuee is 
likely to make his first step outside the cabin;
    (ii) Not less than 0.05 foot-candle (measured normal to the 
direction of the incident light) for a minimum width of 42 inches for a 
Type A overwing emergency exit and two feet for all other overwing 
emergency exits along the 30 percent of the slip-resistant portion of 
the escape route required in Sec. 25.810(c) that is farthest from the 
exit; and
    (iii) Not less than 0.03 foot-candle on the ground surface with the 
landing gear extended (measured normal to the direction of the incident 
light) where an evacuee using the established escape route would 
normally make first contact with the ground.
    (2) At each non-overwing emergency exit not required by Sec. 
25.810(a) to have descent assist means the illumination must be not less 
than 0.03 foot-candle (measured normal to the direction of the incident 
light) on the ground surface with the landing gear extended where an 
evacuee is likely to make first contact with the ground outside the 
cabin.
    (h) The means required in Sec. Sec. 25.810(a)(1) and (d) to assist 
the occupants in descending to the ground must be illuminated so that 
the erected assist means is visible from the airplane.
    (1) If the assist means is illuminated by exterior emergency 
lighting, it must provide illumination of not less than 0.03 foot-candle 
(measured normal to the direction of the incident light) at the ground 
end of the erected assist means where an evacuee using the established 
escape route would normally make first contact with the ground, with the 
airplane in each of the attitudes corresponding to the collapse of one 
or more legs of the landing gear.
    (2) If the emergency lighting subsystem illuminating the assist 
means serves no other assist means, is independent of the airplane's 
main emergency lighting system, and is automatically activated when the 
assist means is erected, the lighting provisions--
    (i) May not be adversely affected by stowage; and
    (ii) Must provide illumination of not less than 0.03 foot-candle 
(measured normal to the direction of incident light) at the ground and 
of the erected assist means where an evacuee would normally make first 
contact with the ground, with the airplane in each of the attitudes 
corresponding to the collapse of one or more legs of the landing gear.
    (i) The energy supply to each emergency lighting unit must provide 
the required level of illumination for at least 10 minutes at the 
critical ambient conditions after emergency landing.
    (j) If storage batteries are used as the energy supply for the 
emergency lighting system, they may be recharged

[[Page 466]]

from the airplane's main electric power system: Provided, That, the 
charging circuit is designed to preclude inadvertent battery discharge 
into charging circuit faults.
    (k) Components of the emergency lighting system, including 
batteries, wiring relays, lamps, and switches must be capable of normal 
operation after having been subjected to the inertia forces listed in 
Sec. 25.561(b).
    (l) The emergency lighting system must be designed so that after any 
single transverse vertical separation of the fuselage during crash 
landing--
    (1) Not more than 25 percent of all electrically illuminated 
emergency lights required by this section are rendered inoperative, in 
addition to the lights that are directly damaged by the separation;
    (2) Each electrically illuminated exit sign required under Sec. 
25.811(d)(2) remains operative exclusive of those that are directly 
damaged by the separation; and
    (3) At least one required exterior emergency light for each side of 
the airplane remains operative exclusive of those that are directly 
damaged by the separation.

[Amdt. 25-15, 32 FR 13265, Sept. 20, 1967, as amended by Amdt. 25-28, 36 
FR 16899, Aug. 26, 1971; Amdt. 25-32, 37 FR 3971, Feb. 24, 1972; Amdt. 
25-46, 43 FR 50597, Oct. 30, 1978; Amdt. 25-58, 49 FR 43186, Oct. 26, 
1984; Amdt. 25-88, 61 FR 57958, Nov. 8, 1996; Amdt. 25-116, 69 FR 62788, 
Oct. 27, 2004; Amdt. 25-128, 74 FR 25645, May 29, 2009]



Sec. 25.813  Emergency exit access.

    Each required emergency exit must be accessible to the passengers 
and located where it will afford an effective means of evacuation. 
Emergency exit distribution must be as uniform as practical, taking 
passenger distribution into account; however, the size and location of 
exits on both sides of the cabin need not be symmetrical. If only one 
floor level exit per side is prescribed, and the airplane does not have 
a tailcone or ventral emergency exit, the floor level exit must be in 
the rearward part of the passenger compartment, unless another location 
affords a more effective means of passenger evacuation. Where more than 
one floor level exit per side is prescribed, at least one floor level 
exit per side must be located near each end of the cabin, except that 
this provision does not apply to combination cargo/passenger 
configurations. In addition--
    (a) There must be a passageway leading from the nearest main aisle 
to each Type A, Type B, Type C, Type I, or Type II emergency exit and 
between individual passenger areas. Each passageway leading to a Type A 
or Type B exit must be unobstructed and at least 36 inches wide. 
Passageways between individual passenger areas and those leading to Type 
I, Type II, or Type C emergency exits must be unobstructed and at least 
20 inches wide. Unless there are two or more main aisles, each Type A or 
B exit must be located so that there is passenger flow along the main 
aisle to that exit from both the forward and aft directions. If two or 
more main aisles are provided, there must be unobstructed cross-aisles 
at least 20 inches wide between main aisles. There must be--
    (1) A cross-aisle which leads directly to each passageway between 
the nearest main aisle and a Type A or B exit; and
    (2) A cross-aisle which leads to the immediate vicinity of each 
passageway between the nearest main aisle and a Type 1, Type II, or Type 
III exit; except that when two Type III exits are located within three 
passenger rows of each other, a single cross-aisle may be used if it 
leads to the vicinity between the passageways from the nearest main 
aisle to each exit.
    (b) Adequate space to allow crewmember(s) to assist in the 
evacuation of passengers must be provided as follows:
    (1) Each assist space must be a rectangle on the floor, of 
sufficient size to enable a crewmember, standing erect, to effectively 
assist evacuees. The assist space must not reduce the unobstructed width 
of the passageway below that required for the exit.
    (2) For each Type A or B exit, assist space must be provided at each 
side of the exit regardless of whether an assist means is required by 
Sec. 25.810(a).
    (3) For each Type C, I or II exit installed in an airplane with 
seating for more than 80 passengers, an assist space must be provided at 
one side of

[[Page 467]]

the passageway regardless of whether an assist means is required by 
Sec. 25.810(a).
    (4) For each Type C, I or II exit, an assist space must be provided 
at one side of the passageway if an assist means is required by Sec. 
25.810(a).
    (5) For any tailcone exit that qualifies for 25 additional passenger 
seats under the provisions of Sec. 25.807(g)(9)(ii), an assist space 
must be provided, if an assist means is required by Sec. 25.810(a).
    (6) There must be a handle, or handles, at each assist space, 
located to enable the crewmember to steady himself or herself:
    (i) While manually activating the assist means (where applicable) 
and,
    (ii) While assisting passengers during an evacuation.
    (c) The following must be provided for each Type III or Type IV 
exit--(1) There must be access from the nearest aisle to each exit. In 
addition, for each Type III exit in an airplane that has a passenger 
seating configuration of 60 or more--
    (i) Except as provided in paragraph (c)(1)(ii), the access must be 
provided by an unobstructed passageway that is at least 10 inches in 
width for interior arrangements in which the adjacent seat rows on the 
exit side of the aisle contain no more than two seats, or 20 inches in 
width for interior arrangements in which those rows contain three seats. 
The width of the passageway must be measured with adjacent seats 
adjusted to their most adverse position. The centerline of the required 
passageway width must not be displaced more than 5 inches horizontally 
from that of the exit.
    (ii) In lieu of one 10- or 20-inch passageway, there may be two 
passageways, between seat rows only, that must be at least 6 inches in 
width and lead to an unobstructed space adjacent to each exit. (Adjacent 
exits must not share a common passageway.) The width of the passageways 
must be measured with adjacent seats adjusted to their most adverse 
position. The unobstructed space adjacent to the exit must extend 
vertically from the floor to the ceiling (or bottom of sidewall stowage 
bins), inboard from the exit for a distance not less than the width of 
the narrowest passenger seat installed on the airplane, and from the 
forward edge of the forward passageway to the aft edge of the aft 
passageway. The exit opening must be totally within the fore and aft 
bounds of the unobstructed space.
    (2) In addition to the access--
    (i) For airplanes that have a passenger seating configuration of 20 
or more, the projected opening of the exit provided must not be 
obstructed and there must be no interference in opening the exit by 
seats, berths, or other protrusions (including any seatback in the most 
adverse position) for a distance from that exit not less than the width 
of the narrowest passenger seat installed on the airplane.
    (ii) For airplanes that have a passenger seating configuration of 19 
or fewer, there may be minor obstructions in this region, if there are 
compensating factors to maintain the effectiveness of the exit.
    (3) For each Type III exit, regardless of the passenger capacity of 
the airplane in which it is installed, there must be placards that--
    (i) Are readable by all persons seated adjacent to and facing a 
passageway to the exit;
    (ii) Accurately state or illustrate the proper method of opening the 
exit, including the use of handholds; and
    (iii) If the exit is a removable hatch, state the weight of the 
hatch and indicate an appropriate location to place the hatch after 
removal.
    (d) If it is necessary to pass through a passageway between 
passenger compartments to reach any required emergency exit from any 
seat in the passenger cabin, the passageway must be unobstructed. 
However, curtains may be used if they allow free entry through the 
passageway.
    (e) No door may be installed between any passenger seat that is 
occupiable for takeoff and landing and any passenger emergency exit, 
such that the door crosses any egress path (including aisles, 
crossaisles and passageways).
    (f) If it is necessary to pass through a doorway separating any 
crewmember seat (except those seats on the flightdeck), occupiable for 
takeoff and landing, from any emergency exit, the door must have a means 
to latch it in the open position. The latching means

[[Page 468]]

must be able to withstand the loads imposed upon it when the door is 
subjected to the ultimate inertia forces, relative to the surrounding 
structure, listed in Sec. 25.561(b).

[Amdt. 25-1, 30 FR 3204, Mar. 9, 1965, as amended by Amdt. 25-15, 32 FR 
13265, Sept. 20, 1967; Amdt. 25-32, 37 FR 3971, Feb. 24, 1972; Amdt. 25-
46, 43 FR 50597, Oct. 30, 1978; Amdt. 25-72, 55 FR 29783, July 20, 1990; 
Amdt. 25-76, 57 FR 19244, May 4, 1992; Amdt. 25-76, 57 FR 29120, June 
30, 1992; Amdt. 25-88, 61 FR 57958, Nov. 8, 1996; Amdt. 25-116, 69 FR 
62788, Oct. 27, 2004; Amdt. 25-128, 74 FR 25645, May 29, 2009]



Sec. 25.815  Width of aisle.

    The passenger aisle width at any point between seats must equal or 
exceed the values in the following table:

------------------------------------------------------------------------
                                                     Minimum passenger
                                                   aisle width (inches)
                                                 -----------------------
           Passenger seating capacity              Less than  25 in. and
                                                    25 in.     more from
                                                  from floor     floor
------------------------------------------------------------------------
10 or less......................................      \1\ 12          15
11 through 19...................................          12          20
20 or more......................................          15          20
------------------------------------------------------------------------
\1\ A narrower width not less than 9 inches may be approved when
  substantiated by tests found necessary by the Administrator.


[Amdt. 25-15, 32 FR 13265, Sept. 20, 1967, as amended by Amdt. 25-38, 41 
FR 55466, Dec. 20, 1976]



Sec. 25.817  Maximum number of seats abreast.

    On airplanes having only one passenger aisle, no more than three 
seats abreast may be placed on each side of the aisle in any one row.

[Amdt. 25-15, 32 FR 13265, Sept. 20, 1967]



Sec. 25.819  Lower deck service compartments (including galleys).

    For airplanes with a service compartment located below the main 
deck, which may be occupied during taxi or flight but not during takeoff 
or landing, the following apply:
    (a) There must be at least two emergency evacuation routes, one at 
each end of each lower deck service compartment or two having sufficient 
separation within each compartment, which could be used by each occupant 
of the lower deck service compartment to rapidly evacuate to the main 
deck under normal and emergency lighting conditions. The routes must 
provide for the evacuation of incapacitated persons, with assistance. 
The use of the evacuation routes may not be dependent on any powered 
device. The routes must be designed to minimize the possibility of 
blockage which might result from fire, mechanical or structural failure, 
or persons standing on top of or against the escape routes. In the event 
the airplane's main power system or compartment main lighting system 
should fail, emergency illumination for each lower deck service 
compartment must be automatically provided.
    (b) There must be a means for two-way voice communication between 
the flight deck and each lower deck service compartment, which remains 
available following loss of normal electrical power generating system.
    (c) There must be an aural emergency alarm system, audible during 
normal and emergency conditions, to enable crewmembers on the flight 
deck and at each required floor level emergency exit to alert occupants 
of each lower deck service compartment of an emergency situation.
    (d) There must be a means, readily detectable by occupants of each 
lower deck service compartment, that indicates when seat belts should be 
fastened.
    (e) If a public address system is installed in the airplane, 
speakers must be provided in each lower deck service compartment.
    (f) For each occupant permitted in a lower deck service compartment, 
there must be a forward or aft facing seat which meets the requirements 
of Sec. 25.785(d), and must be able to withstand maximum flight loads 
when occupied.
    (g) For each powered lift system installed between a lower deck 
service compartment and the main deck for the carriage of persons or 
equipment, or both, the system must meet the following requirements:
    (1) Each lift control switch outside the lift, except emergency stop 
buttons, must be designed to prevent the activation of the life if the 
lift door, or the hatch required by paragraph (g)(3) of this section, or 
both are open.
    (2) An emergency stop button, that when activated will immediately 
stop

[[Page 469]]

the lift, must be installed within the lift and at each entrance to the 
lift.
    (3) There must be a hatch capable of being used for evacuating 
persons from the lift that is openable from inside and outside the lift 
without tools, with the lift in any position.

[Amdt. 25-53, 45 FR 41593, June 19, 1980; 45 FR 43154, June 26, 1980; 
Amdt. 25-110; 68 FR 36883, June 19, 2003]



Sec. 25.820  Lavatory doors.

    All lavatory doors must be designed to preclude anyone from becoming 
trapped inside the lavatory. If a locking mechanism is installed, it 
must be capable of being unlocked from the outside without the aid of 
special tools.

[Doc. No. 2003-14193, 69 FR 24502, May 3, 2004]

                         Ventilation and Heating



Sec. 25.831  Ventilation.

    (a) Under normal operating conditions and in the event of any 
probable failure conditions of any system which would adversely affect 
the ventilating air, the ventilation system must be designed to provide 
a sufficient amount of uncontaminated air to enable the crewmembers to 
perform their duties without undue discomfort or fatigue and to provide 
reasonable passenger comfort. For normal operating conditions, the 
ventilation system must be designed to provide each occupant with an 
airflow containing at least 0.55 pounds of fresh air per minute.
    (b) Crew and passenger compartment air must be free from harmful or 
hazardous concentrations of gases or vapors. In meeting this 
requirement, the following apply:
    (1) Carbon monoxide concentrations in excess of 1 part in 20,000 
parts of air are considered hazardous. For test purposes, any acceptable 
carbon monoxide detection method may be used.
    (2) Carbon dioxide concentration during flight must be shown not to 
exceed 0.5 percent by volume (sea level equivalent) in compartments 
normally occupied by passengers or crewmembers.
    (c) There must be provisions made to ensure that the conditions 
prescribed in paragraph (b) of this section are met after reasonably 
probable failures or malfunctioning of the ventilating, heating, 
pressurization, or other systems and equipment.
    (d) If accumulation of hazardous quantities of smoke in the cockpit 
area is reasonably probable, smoke evacuation must be readily 
accomplished, starting with full pressurization and without 
depressurizing beyond safe limits.
    (e) Except as provided in paragraph (f) of this section, means must 
be provided to enable the occupants of the following compartments and 
areas to control the temperature and quantity of ventilating air 
supplied to their compartment or area independently of the temperature 
and quantity of air supplied to other compartments and areas:
    (1) The flight crew compartment.
    (2) Crewmember compartments and areas other than the flight crew 
compartment unless the crewmember compartment or area is ventilated by 
air interchange with other compartments or areas under all operating 
conditions.
    (f) Means to enable the flight crew to control the temperature and 
quantity of ventilating air supplied to the flight crew compartment 
independently of the temperature and quantity of ventilating air 
supplied to other compartments are not required if all of the following 
conditions are met:
    (1) The total volume of the flight crew and passenger compartments 
is 800 cubic feet or less.
    (2) The air inlets and passages for air to flow between flight crew 
and passenger compartments are arranged to provide compartment 
temperatures within 5 degrees F. of each other and adequate ventilation 
to occupants in both compartments.
    (3) The temperature and ventilation controls are accessible to the 
flight crew.
    (g) The exposure time at any given temperature must not exceed the 
values shown in the following graph after any improbable failure 
condition.

[[Page 470]]

[GRAPHIC] [TIFF OMITTED] TR05JN96.007


[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 
42 FR 36970, July 18, 1977; Amdt. 25-87, 61 FR 28695, June 5, 1996; 
Amdt. 25-89, 61 FR 63956, Dec. 2, 1996]



Sec. 25.832  Cabin ozone concentration.

    (a) The airplane cabin ozone concentration during flight must be 
shown not to exceed--
    (1) 0.25 parts per million by volume, sea level equivalent, at any 
time above flight level 320; and
    (2) 0.1 parts per million by volume, sea level equivalent, time-
weighted average during any 3-hour interval above flight level 270.
    (b) For the purpose of this section, ``sea level equivalent'' refers 
to conditions of 25 [deg]C and 760 millimeters of mercury pressure.
    (c) Compliance with this section must be shown by analysis or tests 
based on airplane operational procedures and performance limitations, 
that demonstrate that either--
    (1) The airplane cannot be operated at an altitude which would 
result in cabin ozone concentrations exceeding the limits prescribed by 
paragraph (a) of this section; or
    (2) The airplane ventilation system, including any ozone control 
equipment, will maintain cabin ozone concentrations at or below the 
limits prescribed by paragraph (a) of this section.

[Amdt. 25-50, 45 FR 3883, Jan. 1, 1980, as amended by Amdt. 25-56, 47 FR 
58489, Dec. 30, 1982; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998]



Sec. 25.833  Combustion heating systems.

    Combustion heaters must be approved.

[Amdt. 25-72, 55 FR 29783, July 20, 1990]

                             Pressurization



Sec. 25.841  Pressurized cabins.

    (a) Pressurized cabins and compartments to be occupied must be 
equipped to provide a cabin pressure altitude of not more than 8,000 
feet at the maximum operating altitude of the airplane under normal 
operating conditions.
    (1) If certification for operation above 25,000 feet is requested, 
the airplane must be designed so that occupants will not be exposed to 
cabin pressure altitudes in excess of 15,000 feet after any probable 
failure condition in the pressurization system.
    (2) The airplane must be designed so that occupants will not be 
exposed to a cabin pressure altitude that exceeds the following after 
decompression from any failure condition not shown to be extremely 
improbable:
    (i) Twenty-five thousand (25,000) feet for more than 2 minutes; or
    (ii) Forty thousand (40,000) feet for any duration.

[[Page 471]]

    (3) Fuselage structure, engine and system failures are to be 
considered in evaluating the cabin decompression.
    (b) Pressurized cabins must have at least the following valves, 
controls, and indicators for controlling cabin pressure:
    (1) Two pressure relief valves to automatically limit the positive 
pressure differential to a predetermined value at the maximum rate of 
flow delivered by the pressure source. The combined capacity of the 
relief valves must be large enough so that the failure of any one valve 
would not cause an appreciable rise in the pressure differential. The 
pressure differential is positive when the internal pressure is greater 
than the external.
    (2) Two reverse pressure differential relief valves (or their 
equivalents) to automatically prevent a negative pressure differential 
that would damage the structure. One valve is enough, however, if it is 
of a design that reasonably precludes its malfunctioning.
    (3) A means by which the pressure differential can be rapidly 
equalized.
    (4) An automatic or manual regulator for controlling the intake or 
exhaust airflow, or both, for maintaining the required internal 
pressures and airflow rates.
    (5) Instruments at the pilot or flight engineer station to show the 
pressure differential, the cabin pressure altitude, and the rate of 
change of the cabin pressure altitude.
    (6) Warning indication at the pilot or flight engineer station to 
indicate when the safe or preset pressure differential and cabin 
pressure altitude limits are exceeded. Appropriate warning markings on 
the cabin pressure differential indicator meet the warning requirement 
for pressure differential limits and an aural or visual signal (in 
addition to cabin altitude indicating means) meets the warning 
requirement for cabin pressure altitude limits if it warns the flight 
crew when the cabin pressure altitude exceeds 10,000 feet.
    (7) A warning placard at the pilot or flight engineer station if the 
structure is not designed for pressure differentials up to the maximum 
relief valve setting in combination with landing loads.
    (8) The pressure sensors necessary to meet the requirements of 
paragraphs (b)(5) and (b)(6) of this section and Sec. 25.1447(c), must 
be located and the sensing system designed so that, in the event of loss 
of cabin pressure in any passenger or crew compartment (including upper 
and lower lobe galleys), the warning and automatic presentation devices, 
required by those provisions, will be actuated without any delay that 
would significantly increase the hazards resulting from decompression.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55466, Dec. 20, 1976; Amdt. 25-87, 61 FR 28696, June 5, 1996]



Sec. 25.843  Tests for pressurized cabins.

    (a) Strength test. The complete pressurized cabin, including doors, 
windows, and valves, must be tested as a pressure vessel for the 
pressure differential specified in Sec. 25.365(d).
    (b) Functional tests. The following functional tests must be 
performed:
    (1) Tests of the functioning and capacity of the positive and 
negative pressure differential valves, and of the emergency release 
valve, to stimulate the effects of closed regulator valves.
    (2) Tests of the pressurization system to show proper functioning 
under each possible condition of pressure, temperature, and moisture, up 
to the maximum altitude for which certification is requested.
    (3) Flight tests, to show the performance of the pressure supply, 
pressure and flow regulators, indicators, and warning signals, in steady 
and stepped climbs and descents at rates corresponding to the maximum 
attainable within the operating limitations of the airplane, up to the 
maximum altitude for which certification is requested.
    (4) Tests of each door and emergency exit, to show that they operate 
properly after being subjected to the flight tests prescribed in 
paragraph (b)(3) of this section.

                             Fire Protection



Sec. 25.851  Fire extinguishers.

    (a) Hand fire extinguishers. (1) The following minimum number of 
hand fire

[[Page 472]]

extinguishers must be conveniently located and evenly distributed in 
passenger compartments:

------------------------------------------------------------------------
               Passenger capacity                  No. of extinguishers
------------------------------------------------------------------------
 7 through 30..................................                        1
 31 through 60.................................                        2
 61 through 200................................                        3
201 through 300................................                        4
301 through 400................................                        5
401 through 500................................                        6
501 through 600................................                        7
601 through 700................................                        8
------------------------------------------------------------------------

    (2) At least one hand fire extinguisher must be conveniently located 
in the pilot compartment.
    (3) At least one readily accessible hand fire extinguisher must be 
available for use in each Class A or Class B cargo or baggage 
compartment and in each Class E cargo or baggage compartment that is 
accessible to crewmembers in flight.
    (4) At least one hand fire extinguisher must be located in, or 
readily accessible for use in, each galley located above or below the 
passenger compartment.
    (5) Each hand fire extinguisher must be approved.
    (6) At least one of the required fire extinguishers located in the 
passenger compartment of an airplane with a passenger capacity of at 
least 31 and not more than 60, and at least two of the fire 
extinguishers located in the passenger compartment of an airplane with a 
passenger capacity of 61 or more must contain Halon 1211 
(bromochlorodifluoromethane CBrC1 F2), or 
equivalent, as the extinguishing agent. The type of extinguishing agent 
used in any other extinguisher required by this section must be 
appropriate for the kinds of fires likely to occur where used.
    (7) The quantity of extinguishing agent used in each extinguisher 
required by this section must be appropriate for the kinds of fires 
likely to occur where used.
    (8) Each extinguisher intended for use in a personnel compartment 
must be designed to minimize the hazard of toxic gas concentration.
    (b) Built-in fire extinguishers. If a built-in fire extinguisher is 
provided--
    (1) Each built-in fire extinguishing system must be installed so 
that--
    (i) No extinguishing agent likely to enter personnel compartments 
will be hazardous to the occupants; and
    (ii) No discharge of the extinguisher can cause structural damage.
    (2) The capacity of each required built-in fire extinguishing system 
must be adequate for any fire likely to occur in the compartment where 
used, considering the volume of the compartment and the ventilation 
rate.

[Amdt. 25-74, 56 FR 15456, Apr. 16, 1991]



Sec. 25.853  Compartment interiors.

    For each compartment occupied by the crew or passengers, the 
following apply:
    (a) Materials (including finishes or decorative surfaces applied to 
the materials) must meet the applicable test criteria prescribed in part 
I of appendix F of this part, or other approved equivalent methods, 
regardless of the passenger capacity of the airplane.
    (b) [Reserved]
    (c) In addition to meeting the requirements of paragraph (a) of this 
section, seat cushions, except those on flight crewmember seats, must 
meet the test requirements of part II of appendix F of this part, or 
other equivalent methods, regardless of the passenger capacity of the 
airplane.
    (d) Except as provided in paragraph (e) of this section, the 
following interior components of airplanes with passenger capacities of 
20 or more must also meet the test requirements of parts IV and V of 
appendix F of this part, or other approved equivalent method, in 
addition to the flammability requirements prescribed in paragraph (a) of 
this section:
    (1) Interior ceiling and wall panels, other than lighting lenses and 
windows;
    (2) Partitions, other than transparent panels needed to enhance 
cabin safety;
    (3) Galley structure, including exposed surfaces of stowed carts and 
standard containers and the cavity walls that are exposed when a full 
complement of such carts or containers is not carried; and
    (4) Large cabinets and cabin stowage compartments, other than 
underseat stowage compartments for stowing small items such as magazines 
and maps.

[[Page 473]]

    (e) The interiors of compartments, such as pilot compartments, 
galleys, lavatories, crew rest quarters, cabinets and stowage 
compartments, need not meet the standards of paragraph (d) of this 
section, provided the interiors of such compartments are isolated from 
the main passenger cabin by doors or equivalent means that would 
normally be closed during an emergency landing condition.
    (f) Smoking is not allowed in lavatories. If smoking is allowed in 
any area occupied by the crew or passengers, an adequate number of self-
contained, removable ashtrays must be provided in designated smoking 
sections for all seated occupants.
    (g) Regardless of whether smoking is allowed in any other part of 
the airplane, lavatories must have self-contained, removable ashtrays 
located conspicuously on or near the entry side of each lavatory door, 
except that one ashtray may serve more than one lavatory door if the 
ashtray can be seen readily from the cabin side of each lavatory served.
    (h) Each receptacle used for the disposal of flammable waste 
material must be fully enclosed, constructed of at least fire resistant 
materials, and must contain fires likely to occur in it under normal 
use. The capability of the receptacle to contain those fires under all 
probable conditions of wear, misalignment, and ventilation expected in 
service must be demonstrated by test.

[Amdt. 25-83, 60 FR 6623, Feb. 2, 1995, as amended by Amdt. 25-116, 69 
FR 62788, Oct. 27, 2004]



Sec. 25.854  Lavatory fire protection.

    For airplanes with a passenger capacity of 20 or more:
    (a) Each lavatory must be equipped with a smoke detector system or 
equivalent that provides a warning light in the cockpit, or provides a 
warning light or audible warning in the passenger cabin that would be 
readily detected by a flight attendant; and
    (b) Each lavatory must be equipped with a built-in fire extinguisher 
for each disposal receptacle for towels, paper, or waste, located within 
the lavatory. The extinguisher must be designed to discharge 
automatically into each disposal receptacle upon occurrence of a fire in 
that receptacle.

[Amdt. 25-74, 56 FR 15456, Apr. 16, 1991]



Sec. 25.855  Cargo or baggage compartments.

    For each cargo or baggage compartment, the following apply:
    (a) The compartment must meet one of the class requirements of Sec. 
25.857.
    (b) Class B through Class E cargo or baggage compartments, as 
defined in Sec. 25.857, must have a liner, and the liner must be 
separate from (but may be attached to) the airplane structure.
    (c) Ceiling and sidewall liner panels of Class C compartments must 
meet the test requirements of part III of appendix F of this part or 
other approved equivalent methods.
    (d) All other materials used in the construction of the cargo or 
baggage compartment must meet the applicable test criteria prescribed in 
part I of appendix F of this part or other approved equivalent methods.
    (e) No compartment may contain any controls, lines, equipment, or 
accessories whose damage or failure would affect safe operation, unless 
those items are protected so that--
    (1) They cannot be damaged by the movement of cargo in the 
compartment, and
    (2) Their breakage or failure will not create a fire hazard.
    (f) There must be means to prevent cargo or baggage from interfering 
with the functioning of the fire protective features of the compartment.
    (g) Sources of heat within the compartment must be shielded and 
insulated to prevent igniting the cargo or baggage.
    (h) Flight tests must be conducted to show compliance with the 
provisions of Sec. 25.857 concerning--
    (1) Compartment accessibility,
    (2) The entries of hazardous quantities of smoke or extinguishing 
agent into compartments occupied by the crew or passengers, and
    (3) The dissipation of the extinguishing agent in Class C 
compartments.
    (i) During the above tests, it must be shown that no inadvertent 
operation of smoke or fire detectors in any compartment would occur as a 
result of

[[Page 474]]

fire contained in any other compartment, either during or after 
extinguishment, unless the extinguishing system floods each such 
compartment simultaneously.
    (j) Cargo or baggage compartment electrical wiring interconnection 
system components must meet the requirements of Sec. 25.1721.

[Amdt. 25-72, 55 FR 29784, July 20, 1990, as amended by Amdt. 25-93, 63 
FR 8048, Feb. 17, 1998; Amdt. 25-116, 69 FR 62788, Oct. 27, 2004; Amdt. 
25-123, 72 FR 63405, Nov. 8, 2007]



Sec. 25.856  Thermal/Acoustic insulation materials.

    (a) Thermal/acoustic insulation material installed in the fuselage 
must meet the flame propagation test requirements of part VI of Appendix 
F to this part, or other approved equivalent test requirements. This 
requirement does not apply to ``small parts,'' as defined in part I of 
Appendix F of this part.
    (b) For airplanes with a passenger capacity of 20 or greater, 
thermal/acoustic insulation materials (including the means of fastening 
the materials to the fuselage) installed in the lower half of the 
airplane fuselage must meet the flame penetration resistance test 
requirements of part VII of Appendix F to this part, or other approved 
equivalent test requirements. This requirement does not apply to 
thermal/acoustic insulation installations that the FAA finds would not 
contribute to fire penetration resistance.

[Amdt. 25-111, 68 FR 45059, July 31, 2003]



Sec. 25.857  Cargo compartment classification.

    (a) Class A; A Class A cargo or baggage compartment is one in 
which--
    (1) The presence of a fire would be easily discovered by a 
crewmember while at his station; and
    (2) Each part of the compartment is easily accessible in flight.
    (b) Class B. A Class B cargo or baggage compartment is one in 
which--
    (1) There is sufficient access in flight to enable a crewmember to 
effectively reach any part of the compartment with the contents of a 
hand fire extinguisher;
    (2) When the access provisions are being used, no hazardous quantity 
of smoke, flames, or extinguishing agent, will enter any compartment 
occupied by the crew or passengers;
    (3) There is a separate approved smoke detector or fire detector 
system to give warning at the pilot or flight engineer station.
    (c) Class C. A Class C cargo or baggage compartment is one not 
meeting the requirements for either a Class A or B compartment but in 
which--
    (1) There is a separate approved smoke detector or fire detector 
system to give warning at the pilot or flight engineer station;
    (2) There is an approved built-in fire extinguishing or suppression 
system controllable from the cockpit.
    (3) There are means to exclude hazardous quantities of smoke, 
flames, or extinguishing agent, from any compartment occupied by the 
crew or passengers;
    (4) There are means to control ventilation and drafts within the 
compartment so that the extinguishing agent used can control any fire 
that may start within the compartment.
    (d) [Reserved]
    (e) Class E. A Class E cargo compartment is one on airplanes used 
only for the carriage of cargo and in which--
    (1) [Reserved]
    (2) There is a separate approved smoke or fire detector system to 
give warning at the pilot or flight engineer station;
    (3) There are means to shut off the ventilating airflow to, or 
within, the compartment, and the controls for these means are accessible 
to the flight crew in the crew compartment;
    (4) There are means to exclude hazardous quantities of smoke, 
flames, or noxious gases, from the flight crew compartment; and
    (5) The required crew emergency exits are accessible under any cargo 
loading condition.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-32, 
37 FR 3972, Feb. 24, 1972; Amdt. 25-60, 51 FR 18243, May 16, 1986; Amdt. 
25-93, 63 FR 8048, Feb. 17, 1998]



Sec. 25.858  Cargo or baggage compartment smoke or fire detection systems.

    If certification with cargo or baggage compartment smoke or fire 
detection

[[Page 475]]

provisions is requested, the following must be met for each cargo or 
baggage compartment with those provisions:
    (a) The detection system must provide a visual indication to the 
flight crew within one minute after the start of a fire.
    (b) The system must be capable of detecting a fire at a temperature 
significantly below that at which the structural integrity of the 
airplane is substantially decreased.
    (c) There must be means to allow the crew to check in flight, the 
functioning of each fire detector circuit.
    (d) The effectiveness of the detection system must be shown for all 
approved operating configurations and conditions.

[Amdt. 25-54, 45 FR 60173, Sept. 11, 1980, as amended by Amdt. 25-93, 63 
FR 8048, Feb. 17, 1998]



Sec. 25.859  Combustion heater fire protection.

    (a) Combustion heater fire zones. The following combustion heater 
fire zones must be protected from fire in accordance with the applicable 
provisions of Sec. Sec. 25.1181 through 25.1191 and Sec. Sec. 25.1195 
through 25.1203;
    (1) The region surrounding the heater, if this region contains any 
flammable fluid system components (excluding the heater fuel system), 
that could--
    (i) Be damaged by heater malfunctioning; or
    (ii) Allow flammable fluids or vapors to reach the heater in case of 
leakage.
    (2) The region surrounding the heater, if the heater fuel system has 
fittings that, if they leaked, would allow fuel or vapors to enter this 
region.
    (3) The part of the ventilating air passage that surrounds the 
combustion chamber. However, no fire extinguishment is required in cabin 
ventilating air passages.
    (b) Ventilating air ducts. Each ventilating air duct passing through 
any fire zone must be fireproof. In addition--
    (1) Unless isolation is provided by fireproof valves or by equally 
effective means, the ventilating air duct downstream of each heater must 
be fireproof for a distance great enough to ensure that any fire 
originating in the heater can be contained in the duct; and
    (2) Each part of any ventilating duct passing through any region 
having a flammable fluid system must be constructed or isolated from 
that system so that the malfunctioning of any component of that system 
cannot introduce flammable fluids or vapors into the ventilating 
airstream.
    (c) Combustion air ducts. Each combustion air duct must be fireproof 
for a distance great enough to prevent damage from backfiring or reverse 
flame propagation. In addition--
    (1) No combustion air duct may have a common opening with the 
ventilating airstream unless flames from backfires or reverse burning 
cannot enter the ventilating airstream under any operating condition, 
including reverse flow or malfunctioning of the heater or its associated 
components; and
    (2) No combustion air duct may restrict the prompt relief of any 
backfire that, if so restricted, could cause heater failure.
    (d) Heater controls; general. Provision must be made to prevent the 
hazardous accumulation of water or ice on or in any heater control 
component, control system tubing, or safety control.
    (e) Heater safety controls. For each combustion heater there must be 
the following safety control means:
    (1) Means independent of the components provided for the normal 
continuous control of air temperature, airflow, and fuel flow must be 
provided, for each heater, to automatically shut off the ignition and 
fuel supply to that heater at a point remote from that heater when any 
of the following occurs:
    (i) The heat exchanger temperature exceeds safe limits.
    (ii) The ventilating air temperature exceeds safe limits.
    (iii) The combustion airflow becomes inadequate for safe operation.
    (iv) The ventilating airflow becomes inadequate for safe operation.
    (2) The means of complying with paragraph (e)(1) of this section for 
any individual heater must--
    (i) Be independent of components serving any other heater whose heat 
output is essential for safe operation; and

[[Page 476]]

    (ii) Keep the heater off until restarted by the crew.
    (3) There must be means to warn the crew when any heater whose heat 
output is essential for safe operation has been shut off by the 
automatic means prescribed in paragraph (e)(1) of this section.
    (f) Air intakes. Each combustion and ventilating air intake must be 
located so that no flammable fluids or vapors can enter the heater 
system under any operating condition--
    (1) During normal operation; or
    (2) As a result of the malfunctioning of any other component.
    (g) Heater exhaust. Heater exhaust systems must meet the provisions 
of Sec. Sec. 25.1121 and 25.1123. In addition, there must be provisions 
in the design of the heater exhaust system to safely expel the products 
of combustion to prevent the occurrence of--
    (1) Fuel leakage from the exhaust to surrounding compartments;
    (2) Exhaust gas impingement on surrounding equipment or structure;
    (3) Ignition of flammable fluids by the exhaust, if the exhaust is 
in a compartment containing flammable fluid lines; and
    (4) Restriction by the exhaust of the prompt relief of backfires 
that, if so restricted, could cause heater failure.
    (h) Heater fuel systems. Each heater fuel system must meet each 
powerplant fuel system requirement affecting safe heater operation. Each 
heater fuel system component within the ventilating airstream must be 
protected by shrouds so that no leakage from those components can enter 
the ventilating airstream.
    (i) Drains. There must be means to safely drain fuel that might 
accumulate within the combustion chamber or the heat exchanger. In 
addition--
    (1) Each part of any drain that operates at high temperatures must 
be protected in the same manner as heater exhausts; and
    (2) Each drain must be protected from hazardous ice accumulation 
under any operating condition.

[Doc. No. 5066, 29 FR 18291, Dec. 24 1964, as amended by Amdt. 25-11, 32 
FR 6912, May 5, 1967; Amdt. 25-23, 35 FR 5676, Apr. 8, 1970]



Sec. 25.863  Flammable fluid fire protection.

    (a) In each area where flammable fluids or vapors might escape by 
leakage of a fluid system, there must be means to minimize the 
probability of ignition of the fluids and vapors, and the resultant 
hazards if ignition does occur.
    (b) Compliance with paragraph (a) of this section must be shown by 
analysis or tests, and the following factors must be considered:
    (1) Possible sources and paths of fluid leakage, and means of 
detecting leakage.
    (2) Flammability characteristics of fluids, including effects of any 
combustible or absorbing materials.
    (3) Possible ignition sources, including electrical faults, 
overheating of equipment, and malfunctioning of protective devices.
    (4) Means available for controlling or extinguishing a fire, such as 
stopping flow of fluids, shutting down equipment, fireproof containment, 
or use of extinguishing agents.
    (5) Ability of airplane components that are critical to safety of 
flight to withstand fire and heat.
    (c) If action by the flight crew is required to prevent or 
counteract a fluid fire (e.g., equipment shutdown or actuation of a fire 
extinguisher) quick acting means must be provided to alert the crew.
    (d) Each area where flammable fluids or vapors might escape by 
leakage of a fluid system must be identified and defined.

[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970, as amended by Amdt. 25-46, 43 FR 
50597, Oct. 30, 1978]



Sec. 25.865  Fire protection of flight controls, engine mounts,
and other flight structure.

    Essential flight controls, engine mounts, and other flight 
structures located in designated fire zones or in adjacent areas which 
would be subjected to the effects of fire in the fire zone must be 
constructed of fireproof material or shielded so that they are capable 
of withstanding the effects of fire.

[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970]

[[Page 477]]



Sec. 25.867  Fire protection: other components.

    (a) Surfaces to the rear of the nacelles, within one nacelle 
diameter of the nacelle centerline, must be at least fire-resistant.
    (b) Paragraph (a) of this section does not apply to tail surfaces to 
the rear of the nacelles that could not be readily affected by heat, 
flames, or sparks coming from a designated fire zone or engine 
compartment of any nacelle.

[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970]



Sec. 25.869  Fire protection: systems.

    (a) Electrical system components:
    (1) Components of the electrical system must meet the applicable 
fire and smoke protection requirements of Sec. Sec. 25.831(c) and 
25.863.
    (2) Equipment that is located in designated fire zones and is used 
during emergency procedures must be at least fire resistant.
    (3) EWIS components must meet the requirements of Sec. 25.1713.
    (b) Each vacuum air system line and fitting on the discharge side of 
the pump that might contain flammable vapors or fluids must meet the 
requirements of Sec. 25.1183 if the line or fitting is in a designated 
fire zone. Other vacuum air systems components in designated fire zones 
must be at least fire resistant.
    (c) Oxygen equipment and lines must--
    (1) Not be located in any designated fire zone,
    (2) Be protected from heat that may be generated in, or escape from, 
any designated fire zone, and
    (3) Be installed so that escaping oxygen cannot cause ignition of 
grease, fluid, or vapor accumulations that are present in normal 
operation or as a result of failure or malfunction of any system.

[Amdt. 25-72, 55 FR 29784, July 20, 1990, as amended by Amdt. 25-113, 69 
FR 12530, Mar. 16, 2004; Amdt. 25-123, 72 FR 63405, Nov. 8, 2007]

                              Miscellaneous



Sec. 25.871  Leveling means.

    There must be means for determining when the airplane is in a level 
position on the ground.

[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970]



Sec. 25.875  Reinforcement near propellers.

    (a) Each part of the airplane near the propeller tips must be strong 
and stiff enough to withstand the effects of the induced vibration and 
of ice thrown from the propeller.
    (b) No window may be near the propeller tips unless it can withstand 
the most severe ice impact likely to occur.



Sec. 25.899  Electrical bonding and protection against static electricity.

    (a) Electrical bonding and protection against static electricity 
must be designed to minimize accumulation of electrostatic charge that 
would cause--
    (1) Human injury from electrical shock,
    (2) Ignition of flammable vapors, or
    (3) Interference with installed electrical/electronic equipment.
    (b) Compliance with paragraph (a) of this section may be shown by--
    (1) Bonding the components properly to the airframe; or
    (2) Incorporating other acceptable means to dissipate the static 
charge so as not to endanger the airplane, personnel, or operation of 
the installed electrical/electronic systems.

[Amdt. 25-123, 72 FR 63405, Nov. 8, 2007]



                          Subpart E_Powerplant

                                 General



Sec. 25.901  Installation.

    (a) For the purpose of this part, the airplane powerplant 
installation includes each component that--
    (1) Is necessary for propulsion;
    (2) Affects the control of the major propulsive units; or
    (3) Affects the safety of the major propulsive units between normal 
inspections or overhauls.
    (b) For each powerplant--
    (1) The installation must comply with--

[[Page 478]]

    (i) The installation instructions provided under Sec. Sec. 33.5 and 
35.3 of this chapter; and
    (ii) The applicable provisions of this subpart;
    (2) The components of the installation must be constructed, 
arranged, and installed so as to ensure their continued safe operation 
between normal inspections or overhauls;
    (3) The installation must be accessible for necessary inspections 
and maintenance; and
    (4) The major components of the installation must be electrically 
bonded to the other parts of the airplane.
    (c) For each powerplant and auxiliary power unit installation, it 
must be established that no single failure or malfunction or probable 
combination of failures will jeopardize the safe operation of the 
airplane except that the failure of structural elements need not be 
considered if the probability of such failure is extremely remote.
    (d) Each auxiliary power unit installation must meet the applicable 
provisions of this subpart.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5676, Apr. 8, 1970; Amdt. 25-40, 42 FR 15042, Mar. 17, 1977; Amdt. 
25-46, 43 FR 50597, Oct. 30, 1978; Amdt. 25-126, 73 FR 63345, Oct. 24, 
2008]



Sec. 25.903  Engines.

    (a) Engine type certificate. (1) Each engine must have a type 
certificate and must meet the applicable requirements of part 34 of this 
chapter.
    (2) Each turbine engine must comply with one of the following:
    (i) Sections 33.76, 33.77 and 33.78 of this chapter in effect on 
December 13, 2000, or as subsequently amended; or
    (ii) Sections 33.77 and 33.78 of this chapter in effect on April 30, 
1998, or as subsequently amended before December 13, 2000; or
    (iii) Comply with Sec. 33.77 of this chapter in effect on October 
31, 1974, or as subsequently amended prior to April 30, 1998, unless 
that engine's foreign object ingestion service history has resulted in 
an unsafe condition; or
    (iv) Be shown to have a foreign object ingestion service history in 
similar installation locations which has not resulted in any unsafe 
condition.

    Note: Sec. 33.77 of this chapter in effect on October 31, 1974, was 
published in 14 CFR parts 1 to 59, Revised as of January 1, 1975. See 39 
FR 35467, October 1, 1974.

    (b) Engine isolation. The powerplants must be arranged and isolated 
from each other to allow operation, in at least one configuration, so 
that the failure or malfunction of any engine, or of any system that can 
affect the engine, will not--
    (1) Prevent the continued safe operation of the remaining engines; 
or
    (2) Require immediate action by any crewmember for continued safe 
operation.
    (c) Control of engine rotation. There must be means for stopping the 
rotation of any engine individually in flight, except that, for turbine 
engine installations, the means for stopping the rotation of any engine 
need be provided only where continued rotation could jeopardize the 
safety of the airplane. Each component of the stopping system on the 
engine side of the firewall that might be exposed to fire must be at 
least fire-resistant. If hydraulic propeller feathering systems are used 
for this purpose, the feathering lines must be at least fire resistant 
under the operating conditions that may be expected to exist during 
feathering.
    (d) Turbine engine installations. For turbine engine installations--
    (1) Design precautions must be taken to minimize the hazards to the 
airplane in the event of an engine rotor failure or of a fire 
originating within the engine which burns through the engine case.
    (2) The powerplant systems associated with engine control devices, 
systems, and instrumentation, must be designed to give reasonable 
assurance that those engine operating limitations that adversely affect 
turbine rotor structural integrity will not be exceeded in service.
    (e) Restart capability. (1) Means to restart any engine in flight 
must be provided.
    (2) An altitude and airspeed envelope must be established for in-
flight engine restarting, and each engine must have a restart capability 
within that envelope.
    (3) For turbine engine powered airplanes, if the minimum windmilling

[[Page 479]]

speed of the engines, following the inflight shutdown of all engines, is 
insufficient to provide the necessary electrical power for engine 
ignition, a power source independent of the engine-driven electrical 
power generating system must be provided to permit in-flight engine 
ignition for restarting.
    (f) Auxiliary Power Unit. Each auxiliary power unit must be approved 
or meet the requirements of the category for its intended use.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5676, Apr. 8, 1970; Amdt. 25-40, 42 FR 15042, Mar. 17, 1977; Amdt. 
25-57, 49 FR 6848, Feb. 23, 1984; Amdt. 25-72, 55 FR 29784, July 20, 
1990; Amdt. 25-73, 55 FR 32861, Aug. 10, 1990; Amdt. 25-94, 63 FR 8848, 
Feb. 23, 1998; Amdt. 25-95, 63 FR 14798, Mar. 26, 1998; Amdt. 25-100, 65 
FR 55854, Sept. 14, 2000]

    Effective Date Note: At 79 FR 65526, Nov. 4, 2014, Sec. 25.903 was 
amended by adding a new paragraph (a)(3), effective Jan. 5, 2015. For 
the convenience of the user, the added text is set forth as follows:



Sec. 25.903  Engines.

    (a) * * *
    (3) Each turbine engine must comply with one of the following 
paragraphs:
    (i) Section 33.68 of this chapter in effect on January 5, 2015, or 
as subsequently amended; or
    (ii) Section 33.68 of this chapter in effect on February 23, 1984, 
or as subsequently amended before January 5, 2015, unless that engine's 
ice accumulation service history has resulted in an unsafe condition; or
    (iii) Section 33.68 of this chapter in effect on October 1, 1974, or 
as subsequently amended prior to February 23, 1984, unless that engine's 
ice accumulation service history has resulted in an unsafe condition; or
    (iv) Be shown to have an ice accumulation service history in similar 
installation locations which has not resulted in any unsafe conditions.

                                * * * * *



Sec. 25.904  Automatic takeoff thrust control system (ATTCS).

    Each applicant seeking approval for installation of an engine power 
control system that automatically resets the power or thrust on the 
operating engine(s) when any engine fails during the takeoff must comply 
with the requirements of appendix I of this part.

[Amdt. 25-62, 52 FR 43156, Nov. 9, 1987]



Sec. 25.905  Propellers.

    (a) Each propeller must have a type certificate.
    (b) Engine power and propeller shaft rotational speed may not exceed 
the limits for which the propeller is certificated.
    (c) The propeller blade pitch control system must meet the 
requirements of Sec. Sec. 35.21, 35.23, 35.42 and 35.43 of this 
chapter.
    (d) Design precautions must be taken to minimize the hazards to the 
airplane in the event a propeller blade fails or is released by a hub 
failure. The hazards which must be considered include damage to 
structure and vital systems due to impact of a failed or released blade 
and the unbalance created by such failure or release.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-54, 
45 FR 60173, Sept. 11, 1980; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984; 
Amdt. 25-72, 55 FR 29784, July 20, 1990; Amdt. 25-126, 73 FR 63345, Oct. 
24, 2008]



Sec. 25.907  Propeller vibration and fatigue.

    This section does not apply to fixed-pitch wood propellers of 
conventional design.
    (a) The applicant must determine the magnitude of the propeller 
vibration stresses or loads, including any stress peaks and resonant 
conditions, throughout the operational envelope of the airplane by 
either:
    (1) Measurement of stresses or loads through direct testing or 
analysis based on direct testing of the propeller on the airplane and 
engine installation for which approval is sought; or
    (2) Comparison of the propeller to similar propellers installed on 
similar airplane installations for which these measurements have been 
made.
    (b) The applicant must demonstrate by tests, analysis based on 
tests, or previous experience on similar designs that the propeller does 
not experience harmful effects of flutter throughout the operational 
envelope of the airplane.
    (c) The applicant must perform an evaluation of the propeller to 
show that failure due to fatigue will be avoided throughout the 
operational life of the propeller using the fatigue and structural data 
obtained in accordance

[[Page 480]]

with part 35 of this chapter and the vibration data obtained from 
compliance with paragraph (a) of this section. For the purpose of this 
paragraph, the propeller includes the hub, blades, blade retention 
component and any other propeller component whose failure due to fatigue 
could be catastrophic to the airplane. This evaluation must include:
    (1) The intended loading spectra including all reasonably 
foreseeable propeller vibration and cyclic load patterns, identified 
emergency conditions, allowable overspeeds and overtorques, and the 
effects of temperatures and humidity expected in service.
    (2) The effects of airplane and propeller operating and 
airworthiness limitations.

[Amdt. 25-126, 73 FR 63345, Oct. 24, 2008]



Sec. 25.925  Propeller clearance.

    Unless smaller clearances are substantiated, propeller clearances 
with the airplane at maximum weight, with the most adverse center of 
gravity, and with the propeller in the most adverse pitch position, may 
not be less than the following:
    (a) Ground clearance. There must be a clearance of at least seven 
inches (for each airplane with nose wheel landing gear) or nine inches 
(for each airplane with tail wheel landing gear) between each propeller 
and the ground with the landing gear statically deflected and in the 
level takeoff, or taxiing attitude, whichever is most critical. In 
addition, there must be positive clearance between the propeller and the 
ground when in the level takeoff attitude with the critical tire(s) 
completely deflated and the corresponding landing gear strut bottomed.
    (b) Water clearance. There must be a clearance of at least 18 inches 
between each propeller and the water, unless compliance with Sec. 
25.239(a) can be shown with a lesser clearance.
    (c) Structural clearance. There must be--
    (1) At least one inch radial clearance between the blade tips and 
the airplane structure, plus any additional radial clearance necessary 
to prevent harmful vibration;
    (2) At least one-half inch longitudinal clearance between the 
propeller blades or cuffs and stationary parts of the airplane; and
    (3) Positive clearance between other rotating parts of the propeller 
or spinner and stationary parts of the airplane.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29784, July 20, 1990]



Sec. 25.929  Propeller deicing.

    (a) For airplanes intended for use where icing may be expected, 
there must be a means to prevent or remove hazardous ice accumulation on 
propellers or on accessories where ice accumulation would jeopardize 
engine performance.
    (b) If combustible fluid is used for propeller deicing, Sec. Sec. 
25.1181 through 25.1185 and 25.1189 apply.

    Effective Date Note: At 79 FR 65526, Nov. 4, 2014, Sec. 25.929 was 
amended by revising paragraph (a), effective Jan. 5, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.929  Propeller deicing.

    (a) If certification for flight in icing is sought there must be a 
means to prevent or remove hazardous ice accumulations that could form 
in the icing conditions defined in Appendix C of this part and in the 
portions of Appendix O of this part for which the airplane is approved 
for flight on propellers or on accessories where ice accumulation would 
jeopardize engine performance.

                                * * * * *



Sec. 25.933  Reversing systems.

    (a) For turbojet reversing systems--
    (1) Each system intended for ground operation only must be designed 
so that during any reversal in flight the engine will produce no more 
than flight idle thrust. In addition, it must be shown by analysis or 
test, or both, that--
    (i) Each operable reverser can be restored to the forward thrust 
position; and
    (ii) The airplane is capable of continued safe flight and landing 
under any possible position of the thrust reverser.
    (2) Each system intended for inflight use must be designed so that 
no unsafe condition will result during normal operation of the system, 
or from any failure (or reasonably likely combination

[[Page 481]]

of failures) of the reversing system, under any anticipated condition of 
operation of the airplane including ground operation. Failure of 
structural elements need not be considered if the probability of this 
kind of failure is extremely remote.
    (3) Each system must have means to prevent the engine from producing 
more than idle thrust when the reversing system malfunctions, except 
that it may produce any greater forward thrust that is shown to allow 
directional control to be maintained, with aerodynamic means alone, 
under the most critical reversing condition expected in operation.
    (b) For propeller reversing systems--
    (1) Each system intended for ground operation only must be designed 
so that no single failure (or reasonably likely combination of failures) 
or malfunction of the system will result in unwanted reverse thrust 
under any expected operating condition. Failure of structural elements 
need not be considered if this kind of failure is extremely remote.
    (2) Compliance with this section may be shown by failure analysis or 
testing, or both, for propeller systems that allow propeller blades to 
move from the flight low-pitch position to a position that is 
substantially less than that at the normal flight low-pitch position. 
The analysis may include or be supported by the analysis made to show 
compliance with the requirements of Sec. 35.21 of this chapter for the 
propeller and associated installation components.

[Amdt. 25-72, 55 FR 29784, July 20, 1990]



Sec. 25.934  Turbojet engine thrust reverser system tests.

    Thrust reversers installed on turbojet engines must meet the 
requirements of Sec. 33.97 of this chapter.

[Amdt. 25-23, 35 FR 5677, Apr. 8, 1970]



Sec. 25.937  Turbopropeller-drag limiting systems.

    Turbopropeller power airplane propeller-drag limiting systems must 
be designed so that no single failure or malfunction of any of the 
systems during normal or emergency operation results in propeller drag 
in excess of that for which the airplane was designed under Sec. 
25.367. Failure of structural elements of the drag limiting systems need 
not be considered if the probability of this kind of failure is 
extremely remote.



Sec. 25.939  Turbine engine operating characteristics.

    (a) Turbine engine operating characteristics must be investigated in 
flight to determine that no adverse characteristics (such as stall, 
surge, or flameout) are present, to a hazardous degree, during normal 
and emergency operation within the range of operating limitations of the 
airplane and of the engine.
    (b) [Reserved]
    (c) The turbine engine air inlet system may not, as a result of air 
flow distortion during normal operation, cause vibration harmful to the 
engine.

[Amdt. 25-11, 32 FR 6912, May 5, 1967, as amended by Amdt. 25-40, 42 FR 
15043, Mar. 17, 1977]



Sec. 25.941  Inlet, engine, and exhaust compatibility.

    For airplanes using variable inlet or exhaust system geometry, or 
both--
    (a) The system comprised of the inlet, engine (including thrust 
augmentation systems, if incorporated), and exhaust must be shown to 
function properly under all operating conditions for which approval is 
sought, including all engine rotating speeds and power settings, and 
engine inlet and exhaust configurations;
    (b) The dynamic effects of the operation of these (including 
consideration of probable malfunctions) upon the aerodynamic control of 
the airplane may not result in any condition that would require 
exceptional skill, alertness, or strength on the part of the pilot to 
avoid exceeding an operational or structural limitation of the airplane; 
and
    (c) In showing compliance with paragraph (b) of this section, the 
pilot strength required may not exceed the limits set forth in Sec. 
25.143(d), subject to the conditions set forth in paragraphs (e) and (f) 
of Sec. 25.143.

[Amdt. 25-38, 41 FR 55467, Dec. 20, 1976, as amended by Amdt. 25-121, 72 
FR 44669, Aug. 8, 2007]

[[Page 482]]



Sec. 25.943  Negative acceleration.

    No hazardous malfunction of an engine, an auxiliary power unit 
approved for use in flight, or any component or system associated with 
the powerplant or auxiliary power unit may occur when the airplane is 
operated at the negative accelerations within the flight envelopes 
prescribed in Sec. 25.333. This must be shown for the greatest duration 
expected for the acceleration.

[Amdt. 25-40, 42 FR 15043, Mar. 17, 1977]



Sec. 25.945  Thrust or power augmentation system.

    (a) General. Each fluid injection system must provide a flow of 
fluid at the rate and pressure established for proper engine functioning 
under each intended operating condition. If the fluid can freeze, fluid 
freezing may not damage the airplane or adversely affect airplane 
performance.
    (b) Fluid tanks. Each augmentation system fluid tank must meet the 
following requirements:
    (1) Each tank must be able to withstand without failure the 
vibration, inertia, fluid, and structural loads that it may be subject 
to in operation.
    (2) The tanks as mounted in the airplane must be able to withstand 
without failure or leakage an internal pressure 1.5 times the maximum 
operating pressure.
    (3) If a vent is provided, the venting must be effective under all 
normal flight conditions.
    (4) [Reserved]
    (5) Each tank must have an expansion space of not less than 2 
percent of the tank capacity. It must be impossible to fill the 
expansion space inadvertently with the airplane in the normal ground 
attitude.
    (c) Augmentation system drains must be designed and located in 
accordance with Sec. 25.1455 if--
    (1) The augmentation system fluid is subject to freezing; and
    (2) The fluid may be drained in flight or during ground operation.
    (d) The augmentation liquid tank capacity available for the use of 
each engine must be large enough to allow operation of the airplane 
under the approved procedures for the use of liquid-augmented power. The 
computation of liquid consumption must be based on the maximum approved 
rate appropriate for the desired engine output and must include the 
effect of temperature on engine performance as well as any other factors 
that might vary the amount of liquid required.
    (e) This section does not apply to fuel injection systems.

[Amdt. 25-40, 42 FR 15043, Mar. 17, 1977, as amended by Amdt. 25-72, 55 
FR 29785, July 20, 1990; Amdt. 25-115, 69 FR 40527, July 2, 2004]

                               Fuel System



Sec. 25.951  General.

    (a) Each fuel system must be constructed and arranged to ensure a 
flow of fuel at a rate and pressure established for proper engine and 
auxiliary power unit functioning under each likely operating condition, 
including any maneuver for which certification is requested and during 
which the engine or auxiliary power unit is permitted to be in 
operation.
    (b) Each fuel system must be arranged so that any air which is 
introduced into the system will not result in--
    (1) Power interruption for more than 20 seconds for reciprocating 
engines; or
    (2) Flameout for turbine engines.
    (c) Each fuel system for a turbine engine must be capable of 
sustained operation throughout its flow and pressure range with fuel 
initially saturated with water at 80 [deg]F and having 0.75cc of free 
water per gallon added and cooled to the most critical condition for 
icing likely to be encountered in operation.
    (d) Each fuel system for a turbine engine powered airplane must meet 
the applicable fuel venting requirements of part 34 of this chapter.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5677, Apr. 8, 1970; Amdt. 25-36, 39 FR 35460, Oct. 1, 1974; Amdt. 
25-38, 41 FR 55467, Dec. 20, 1976; Amdt. 25-73, 55 FR 32861, Aug. 10, 
1990]



Sec. 25.952  Fuel system analysis and test.

    (a) Proper fuel system functioning under all probable operating 
conditions must be shown by analysis and those tests found necessary by 
the Administrator. Tests, if required, must be made using the airplane 
fuel system or a test article that reproduces the operating

[[Page 483]]

characteristics of the portion of the fuel system to be tested.
    (b) The likely failure of any heat exchanger using fuel as one of 
its fluids may not result in a hazardous condition.

[Amdt. 25-40, 42 FR 15043, Mar. 17, 1977]



Sec. 25.953  Fuel system independence.

    Each fuel system must meet the requirements of Sec. 25.903(b) by--
    (a) Allowing the supply of fuel to each engine through a system 
independent of each part of the system supplying fuel to any other 
engine; or
    (b) Any other acceptable method.



Sec. 25.954  Fuel system lightning protection.

    The fuel system must be designed and arranged to prevent the 
ignition of fuel vapor within the system by--
    (a) Direct lightning strikes to areas having a high probability of 
stroke attachment;
    (b) Swept lightning strokes to areas where swept strokes are highly 
probable; and
    (c) Corona and streamering at fuel vent outlets.

[Amdt. 25-14, 32 FR 11629, Aug. 11, 1967]



Sec. 25.955  Fuel flow.

    (a) Each fuel system must provide at least 100 percent of the fuel 
flow required under each intended operating condition and maneuver. 
Compliance must be shown as follows:
    (1) Fuel must be delivered to each engine at a pressure within the 
limits specified in the engine type certificate.
    (2) The quantity of fuel in the tank may not exceed the amount 
established as the unusable fuel supply for that tank under the 
requirements of Sec. 25.959 plus that necessary to show compliance with 
this section.
    (3) Each main pump must be used that is necessary for each operating 
condition and attitude for which compliance with this section is shown, 
and the appropriate emergency pump must be substituted for each main 
pump so used.
    (4) If there is a fuel flowmeter, it must be blocked and the fuel 
must flow through the meter or its bypass.
    (b) If an engine can be supplied with fuel from more than one tank, 
the fuel system must--
    (1) For each reciprocating engine, supply the full fuel pressure to 
that engine in not more than 20 seconds after switching to any other 
fuel tank containing usable fuel when engine malfunctioning becomes 
apparent due to the depletion of the fuel supply in any tank from which 
the engine can be fed; and
    (2) For each turbine engine, in addition to having appropriate 
manual switching capability, be designed to prevent interruption of fuel 
flow to that engine, without attention by the flight crew, when any tank 
supplying fuel to that engine is depleted of usable fuel during normal 
operation, and any other tank, that normally supplies fuel to that 
engine alone, contains usable fuel.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 
32 FR 6912, May 5, 1967]



Sec. 25.957  Flow between interconnected tanks.

    If fuel can be pumped from one tank to another in flight, the fuel 
tank vents and the fuel transfer system must be designed so that no 
structural damage to the tanks can occur because of overfilling.



Sec. 25.959  Unusable fuel supply.

    The unusable fuel quantity for each fuel tank and its fuel system 
components must be established at not less than the quantity at which 
the first evidence of engine malfunction occurs under the most adverse 
fuel feed condition for all intended operations and flight maneuvers 
involving fuel feeding from that tank. Fuel system component failures 
need not be considered.

[Amdt. 25-23, 35 FR 5677, Apr. 8, 1970, as amended by Amdt. 25-40, 42 FR 
15043, Mar. 17, 1977]



Sec. 25.961  Fuel system hot weather operation.

    (a) The fuel system must perform satisfactorily in hot weather 
operation. This must be shown by showing that the fuel system from the 
tank outlets to each engine is pressurized, under all intended 
operations, so as to prevent

[[Page 484]]

vapor formation, or must be shown by climbing from the altitude of the 
airport elected by the applicant to the maximum altitude established as 
an operating limitation under Sec. 25.1527. If a climb test is elected, 
there may be no evidence of vapor lock or other malfunctioning during 
the climb test conducted under the following conditions:
    (1) For reciprocating engine powered airplanes, the engines must 
operate at maximum continuous power, except that takeoff power must be 
used for the altitudes from 1,000 feet below the critical altitude 
through the critical altitude. The time interval during which takeoff 
power is used may not be less than the takeoff time limitation.
    (2) For turbine engine powered airplanes, the engines must operate 
at takeoff power for the time interval selected for showing the takeoff 
flight path, and at maximum continuous power for the rest of the climb.
    (3) The weight of the airplane must be the weight with full fuel 
tanks, minimum crew, and the ballast necessary to maintain the center of 
gravity within allowable limits.
    (4) The climb airspeed may not exceed--
    (i) For reciprocating engine powered airplanes, the maximum airspeed 
established for climbing from takeoff to the maximum operating altitude 
with the airplane in the following configuration:
    (A) Landing gear retracted.
    (B) Wing flaps in the most favorable position.
    (C) Cowl flaps (or other means of controlling the engine cooling 
supply) in the position that provides adequate cooling in the hot-day 
condition.
    (D) Engine operating within the maximum continuous power 
limitations.
    (E) Maximum takeoff weight; and
    (ii) For turbine engine powered airplanes, the maximum airspeed 
established for climbing from takeoff to the maximum operating altitude.
    (5) The fuel temperature must be at least 110 [deg]F.
    (b) The test prescribed in paragraph (a) of this section may be 
performed in flight or on the ground under closely simulated flight 
conditions. If a flight test is performed in weather cold enough to 
interfere with the proper conduct of the test, the fuel tank surfaces, 
fuel lines, and other fuel system parts subject to cold air must be 
insulated to simulate, insofar as practicable, flight in hot weather.

[Amdt. 25-11, 32 FR 6912, May 5, 1967, as amended by Amdt. 25-57, 49 FR 
6848, Feb. 23, 1984]



Sec. 25.963  Fuel tanks: general.

    (a) Each fuel tank must be able to withstand, without failure, the 
vibration, inertia, fluid, and structural loads that it may be subjected 
to in operation.
    (b) Flexible fuel tank liners must be approved or must be shown to 
be suitable for the particular application.
    (c) Integral fuel tanks must have facilities for interior inspection 
and repair.
    (d) Fuel tanks must, so far as it is practicable, be designed, 
located, and installed so that no fuel is released in or near the 
fuselage, or near the engines, in quantities that would constitute a 
fire hazard in otherwise survivable emergency landing conditions, and--
    (1) Fuel tanks must be able to resist rupture and retain fuel under 
ultimate hydrostatic design conditions in which the pressure P within 
the tank varies in accordance with the formula:

P = K[rho]gL

Where--

P = fuel pressure at each point within the tank
[rho] = typical fuel density
g = acceleration due to gravity
L = a reference distance between the point of pressure and the tank 
          farthest boundary in the direction of loading
K = 4.5 for the forward loading condition for those parts of fuel tanks 
          outside the fuselage pressure boundary
K = 9 for the forward loading condition for those parts of fuel tanks 
          within the fuselage pressure boundary, or that form part of 
          the fuselage pressure boundary
K = 1.5 for the aft loading condition
K = 3.0 for the inboard and outboard loading conditions for those parts 
          of fuel tanks within the fuselage pressure boundary, or that 
          form part of the fuselage pressure boundary
K = 1.5 for the inboard and outboard loading conditions for those parts 
          of fuel tanks outside the fuselage pressure boundary
K = 6 for the downward loading condition
K = 3 for the upward loading condition


[[Page 485]]


    (2) For those parts of wing fuel tanks near the fuselage or near the 
engines, the greater of the fuel pressures resulting from paragraphs 
(d)(2)(i) or (d)(2)(ii) of this section must be used:
    (i) The fuel pressures resulting from paragraph (d)(1) of this 
section, and
    (ii) The lesser of the two following conditions:
    (A) Fuel pressures resulting from the accelerations specified in 
Sec. 25.561(b)(3) considering the fuel tank full of fuel at maximum 
fuel density. Fuel pressures based on the 9.0g forward acceleration may 
be calculated using the fuel static head equal to the streamwise local 
chord of the tank. For inboard and outboard conditions, an acceleration 
of 1.5g may be used in lieu of 3.0g as specified in Sec. 25.561(b)(3).
    (B) Fuel pressures resulting from the accelerations as specified in 
Sec. 25.561(b)(3) considering a fuel volume beyond 85 percent of the 
maximum permissible volume in each tank using the static head associated 
with the 85 percent fuel level. A typical density of the appropriate 
fuel may be used. For inboard and outboard conditions, an acceleration 
of 1.5g may be used in lieu of 3.0g as specified in Sec. 25.561(b)(3).
    (3) Fuel tank internal barriers and baffles may be considered as 
solid boundaries if shown to be effective in limiting fuel flow.
    (4) For each fuel tank and surrounding airframe structure, the 
effects of crushing and scraping actions with the ground must not cause 
the spillage of enough fuel, or generate temperatures that would 
constitute a fire hazard under the conditions specified in Sec. 
25.721(b).
    (5) Fuel tank installations must be such that the tanks will not 
rupture as a result of the landing gear or an engine pylon or engine 
mount tearing away as specified in Sec. 25.721(a) and (c).
    (e) Fuel tank access covers must comply with the following criteria 
in order to avoid loss of hazardous quantities of fuel:
    (1) All covers located in an area where experience or analysis 
indicates a strike is likely must be shown by analysis or tests to 
minimize penetration and deformation by tire fragments, low energy 
engine debris, or other likely debris.
    (2) All covers must be fire resistant as defined in part 1 of this 
chapter.
    (f) For pressurized fuel tanks, a means with fail-safe features must 
be provided to prevent the buildup of an excessive pressure difference 
between the inside and the outside of the tank.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 
42 FR 15043, Mar. 17, 1977; Amdt. 25-69, 54 FR 40354, Sept. 29, 1989; 
Amdt. No. 25-139, 79 FR 59430, Oct. 2, 2014]



Sec. 25.965  Fuel tank tests.

    (a) It must be shown by tests that the fuel tanks, as mounted in the 
airplane, can withstand, without failure or leakage, the more critical 
of the pressures resulting from the conditions specified in paragraphs 
(a)(1) and (2) of this section. In addition, it must be shown by either 
analysis or tests, that tank surfaces subjected to more critical 
pressures resulting from the condition of paragraphs (a)(3) and (4) of 
this section, are able to withstand the following pressures:
    (1) An internal pressure of 3.5 psi.
    (2) 125 percent of the maximum air pressure developed in the tank 
from ram effect.
    (3) Fluid pressures developed during maximum limit accelerations, 
and deflections, of the airplane with a full tank.
    (4) Fluid pressures developed during the most adverse combination of 
airplane roll and fuel load.
    (b) Each metallic tank with large unsupported or unstiffened flat 
surfaces, whose failure or deformation could cause fuel leakage, must be 
able to withstand the following test, or its equivalent, without leakage 
or excessive deformation of the tank walls:
    (1) Each complete tank assembly and its supports must be vibration 
tested while mounted to simulate the actual installation.
    (2) Except as specified in paragraph (b)(4) of this section, the 
tank assembly must be vibrated for 25 hours at an amplitude of not less 
than \1/32\ of an inch (unless another amplitude is substantiated) while 
\2/3\ filled with water or other suitable test fluid.
    (3) The test frequency of vibration must be as follows:
    (i) If no frequency of vibration resulting from any r.p.m. within 
the normal

[[Page 486]]

operating range of engine speeds is critical, the test frequency of 
vibration must be 2,000 cycles per minute.
    (ii) If only one frequency of vibration resulting from any r.p.m. 
within the normal operating range of engine speeds is critical, that 
frequency of vibration must be the test frequency.
    (iii) If more than one frequency of vibration resulting from any 
r.p.m. within the normal operating range of engine speeds is critical, 
the most critical of these frequencies must be the test frequency.
    (4) Under paragraphs (b)(3)(ii) and (iii) of this section, the time 
of test must be adjusted to accomplish the same number of vibration 
cycles that would be accomplished in 25 hours at the frequency specified 
in paragraph (b)(3)(i) of this section.
    (5) During the test, the tank assembly must be rocked at the rate of 
16 to 20 complete cycles per minute, through an angle of 15[deg] on both 
sides of the horizontal (30[deg] total), about the most critical axis, 
for 25 hours. If motion about more than one axis is likely to be 
critical, the tank must be rocked about each critical axis for 12\1/2\ 
hours.
    (c) Except where satisfactory operating experience with a similar 
tank in a similar installation is shown, nonmetallic tanks must 
withstand the test specified in paragraph (b)(5) of this section, with 
fuel at a temperature of 110 [deg]F. During this test, a representative 
specimen of the tank must be installed in a supporting structure 
simulating the installation in the airplane.
    (d) For pressurized fuel tanks, it must be shown by analysis or 
tests that the fuel tanks can withstand the maximum pressure likely to 
occur on the ground or in flight.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 
32 FR 6913, May 5, 1967; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977]



Sec. 25.967  Fuel tank installations.

    (a) Each fuel tank must be supported so that tank loads (resulting 
from the weight of the fuel in the tanks) are not concentrated on 
unsupported tank surfaces. In addition--
    (1) There must be pads, if necessary, to prevent chafing between the 
tank and its supports;
    (2) Padding must be nonabsorbent or treated to prevent the 
absorption of fluids;
    (3) If a flexible tank liner is used, it must be supported so that 
it is not required to withstand fluid loads; and
    (4) Each interior surface of the tank compartment must be smooth and 
free of projections that could cause wear of the liner unless--
    (i) Provisions are made for protection of the liner at these points; 
or
    (ii) The construction of the liner itself provides that protection.
    (b) Spaces adjacent to tank surfaces must be ventilated to avoid 
fume accumulation due to minor leakage. If the tank is in a sealed 
compartment, ventilation may be limited to drain holes large enough to 
prevent excessive pressure resulting from altitude changes.
    (c) The location of each tank must meet the requirements of Sec. 
25.1185(a).
    (d) No engine nacelle skin immediately behind a major air outlet 
from the engine compartment may act as the wall of an integral tank.
    (e) Each fuel tank must be isolated from personnel compartments by a 
fumeproof and fuelproof enclosure.



Sec. 25.969  Fuel tank expansion space.

    Each fuel tank must have an expansion space of not less than 2 
percent of the tank capacity. It must be impossible to fill the 
expansion space inadvertently with the airplane in the normal ground 
attitude. For pressure fueling systems, compliance with this section may 
be shown with the means provided to comply with Sec. 25.979(b).

[Amdt. 25-11, 32 FR 6913, May 5, 1967]



Sec. 25.971  Fuel tank sump.

    (a) Each fuel tank must have a sump with an effective capacity, in 
the normal ground attitude, of not less than the greater of 0.10 percent 
of the tank capacity or one-sixteenth of a gallon unless operating 
limitations are established to ensure that the accumulation of water in 
service will not exceed the sump capacity.
    (b) Each fuel tank must allow drainage of any hazardous quantity of 
water from any part of the tank to its sump with the airplane in the 
ground attitude.

[[Page 487]]

    (c) Each fuel tank sump must have an accessible drain that--
    (1) Allows complete drainage of the sump on the ground;
    (2) Discharges clear of each part of the airplane; and
    (3) Has manual or automatic means for positive locking in the closed 
position.



Sec. 25.973  Fuel tank filler connection.

    Each fuel tank filler connection must prevent the entrance of fuel 
into any part of the airplane other than the tank itself. In addition--
    (a) [Reserved]
    (b) Each recessed filler connection that can retain any appreciable 
quantity of fuel must have a drain that discharges clear of each part of 
the airplane;
    (c) Each filler cap must provide a fuel-tight seal; and
    (d) Each fuel filling point must have a provision for electrically 
bonding the airplane to ground fueling equipment.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 
42 FR 15043, Mar. 17, 1977; Amdt. 25-72, 55 FR 29785, July 20, 1990; 
Amdt. 25-115, 69 FR 40527, July 2, 2004]



Sec. 25.975  Fuel tank vents and carburetor vapor vents.

    (a) Fuel tank vents. Each fuel tank must be vented from the top part 
of the expansion space so that venting is effective under any normal 
flight condition. In addition--
    (1) Each vent must be arranged to avoid stoppage by dirt or ice 
formation;
    (2) The vent arrangement must prevent siphoning of fuel during 
normal operation;
    (3) The venting capacity and vent pressure levels must maintain 
acceptable differences of pressure between the interior and exterior of 
the tank, during--
    (i) Normal flight operation;
    (ii) Maximum rate of ascent and descent; and
    (iii) Refueling and defueling (where applicable);
    (4) Airspaces of tanks with interconnected outlets must be 
interconnected;
    (5) There may be no point in any vent line where moisture can 
accumulate with the airplane in the ground attitude or the level flight 
attitude, unless drainage is provided; and
    (6) No vent or drainage provision may end at any point--
    (i) Where the discharge of fuel from the vent outlet would 
constitute a fire hazard; or
    (ii) From which fumes could enter personnel compartments.
    (b) Carburetor vapor vents. Each carburetor with vapor elimination 
connections must have a vent line to lead vapors back to one of the fuel 
tanks. In addition--
    (1) Each vent system must have means to avoid stoppage by ice; and
    (2) If there is more than one fuel tank, and it is necessary to use 
the tanks in a definite sequence, each vapor vent return line must lead 
back to the fuel tank used for takeoff and landing.



Sec. 25.977  Fuel tank outlet.

    (a) There must be a fuel strainer for the fuel tank outlet or for 
the booster pump. This strainer must--
    (1) For reciprocating engine powered airplanes, have 8 to 16 meshes 
per inch; and
    (2) For turbine engine powered airplanes, prevent the passage of any 
object that could restrict fuel flow or damage any fuel system 
component.
    (b) [Reserved]
    (c) The clear area of each fuel tank outlet strainer must be at 
least five times the area of the outlet line.
    (d) The diameter of each strainer must be at least that of the fuel 
tank outlet.
    (e) Each finger strainer must be accessible for inspection and 
cleaning.

[Amdt. 25-11, 32 FR 6913, May 5, 1967, as amended by Amdt. 25-36, 39 FR 
35460, Oct. 1, 1974]



Sec. 25.979  Pressure fueling system.

    For pressure fueling systems, the following apply:
    (a) Each pressure fueling system fuel manifold connection must have 
means to prevent the escape of hazardous quantities of fuel from the 
system if the fuel entry valve fails.
    (b) An automatic shutoff means must be provided to prevent the 
quantity of fuel in each tank from exceeding the

[[Page 488]]

maximum quantity approved for that tank. This means must--
    (1) Allow checking for proper shutoff operation before each fueling 
of the tank; and
    (2) Provide indication at each fueling station of failure of the 
shutoff means to stop the fuel flow at the maximum quantity approved for 
that tank.
    (c) A means must be provided to prevent damage to the fuel system in 
the event of failure of the automatic shutoff means prescribed in 
paragraph (b) of this section.
    (d) The airplane pressure fueling system (not including fuel tanks 
and fuel tank vents) must withstand an ultimate load that is 2.0 times 
the load arising from the maximum pressures, including surge, that is 
likely to occur during fueling. The maximum surge pressure must be 
established with any combination of tank valves being either 
intentionally or inadvertently closed.
    (e) The airplane defueling system (not including fuel tanks and fuel 
tank vents) must withstand an ultimate load that is 2.0 times the load 
arising from the maximum permissible defueling pressure (positive or 
negative) at the airplane fueling connection.

[Amdt. 25-11, 32 FR 6913, May 5, 1967, as amended by Amdt. 25-38, 41 FR 
55467, Dec. 20, 1976; Amdt. 25-72, 55 FR 29785, July 20, 1990]



Sec. 25.981  Fuel tank ignition prevention.

    (a) No ignition source may be present at each point in the fuel tank 
or fuel tank system where catastrophic failure could occur due to 
ignition of fuel or vapors. This must be shown by:
    (1) Determining the highest temperature allowing a safe margin below 
the lowest expected autoignition temperature of the fuel in the fuel 
tanks.
    (2) Demonstrating that no temperature at each place inside each fuel 
tank where fuel ignition is possible will exceed the temperature 
determined under paragraph (a)(1) of this section. This must be verified 
under all probable operating, failure, and malfunction conditions of 
each component whose operation, failure, or malfunction could increase 
the temperature inside the tank.
    (3) Demonstrating that an ignition source could not result from each 
single failure, from each single failure in combination with each latent 
failure condition not shown to be extremely remote, and from all 
combinations of failures not shown to be extremely improbable. The 
effects of manufacturing variability, aging, wear, corrosion, and likely 
damage must be considered.
    (b) Except as provided in paragraphs (b)(2) and (c) of this section, 
no fuel tank Fleet Average Flammability Exposure on an airplane may 
exceed three percent of the Flammability Exposure Evaluation Time (FEET) 
as defined in Appendix N of this part, or that of a fuel tank within the 
wing of the airplane model being evaluated, whichever is greater. If the 
wing is not a conventional unheated aluminum wing, the analysis must be 
based on an assumed Equivalent Conventional Unheated Aluminum Wing Tank.
    (1) Fleet Average Flammability Exposure is determined in accordance 
with Appendix N of this part. The assessment must be done in accordance 
with the methods and procedures set forth in the Fuel Tank Flammability 
Assessment Method User's Manual, dated May 2008, document number DOT/
FAA/AR-05/8 (incorporated by reference, see Sec. 25.5).
    (2) Any fuel tank other than a main fuel tank on an airplane must 
meet the flammability exposure criteria of Appendix M to this part if 
any portion of the tank is located within the fuselage contour.
    (3) As used in this paragraph,
    (i) Equivalent Conventional Unheated Aluminum Wing Tank is an 
integral tank in an unheated semi-monocoque aluminum wing of a subsonic 
airplane that is equivalent in aerodynamic performance, structural 
capability, fuel tank capacity and tank configuration to the designed 
wing.
    (ii) Fleet Average Flammability Exposure is defined in Appendix N to 
this part and means the percentage of time each fuel tank ullage is 
flammable for a fleet of an airplane type operating over the range of 
flight lengths.
    (iii) Main Fuel Tank means a fuel tank that feeds fuel directly into 
one or more engines and holds required fuel reserves continually 
throughout each flight.

[[Page 489]]

    (c) Paragraph (b) of this section does not apply to a fuel tank if 
means are provided to mitigate the effects of an ignition of fuel vapors 
within that fuel tank such that no damage caused by an ignition will 
prevent continued safe flight and landing.
    (d) Critical design configuration control limitations (CDCCL), 
inspections, or other procedures must be established, as necessary, to 
prevent development of ignition sources within the fuel tank system 
pursuant to paragraph (a) of this section, to prevent increasing the 
flammability exposure of the tanks above that permitted under paragraph 
(b) of this section, and to prevent degradation of the performance and 
reliability of any means provided according to paragraphs (a) or (c) of 
this section. These CDCCL, inspections, and procedures must be included 
in the Airworthiness Limitations section of the instructions for 
continued airworthiness required by Sec. 25.1529. Visible means of 
identifying critical features of the design must be placed in areas of 
the airplane where foreseeable maintenance actions, repairs, or 
alterations may compromise the critical design configuration control 
limitations (e.g., color-coding of wire to identify separation 
limitation). These visible means must also be identified as CDCCL.

[Doc. No. 1999-6411, 66 FR 23129, May 7, 2001, as amended by Docket No. 
FAA-2005-22997, 73 FR 42494, July 21, 2008]

                         Fuel System Components



Sec. 25.991  Fuel pumps.

    (a) Main pumps. Each fuel pump required for proper engine operation, 
or required to meet the fuel system requirements of this subpart (other 
than those in paragraph (b) of this section, is a main pump. For each 
main pump, provision must be made to allow the bypass of each positive 
displacement fuel pump other than a fuel injection pump (a pump that 
supplies the proper flow and pressure for fuel injection when the 
injection is not accomplished in a carburetor) approved as part of the 
engine.
    (b) Emergency pumps. There must be emergency pumps or another main 
pump to feed each engine immediately after failure of any main pump 
(other than a fuel injection pump approved as part of the engine).



Sec. 25.993  Fuel system lines and fittings.

    (a) Each fuel line must be installed and supported to prevent 
excessive vibration and to withstand loads due to fuel pressure and 
accelerated flight conditions.
    (b) Each fuel line connected to components of the airplane between 
which relative motion could exist must have provisions for flexibility.
    (c) Each flexible connection in fuel lines that may be under 
pressure and subjected to axial loading must use flexible hose 
assemblies.
    (d) Flexible hose must be approved or must be shown to be suitable 
for the particular application.
    (e) No flexible hose that might be adversely affected by exposure to 
high temperatures may be used where excessive temperatures will exist 
during operation or after engine shut-down.
    (f) Each fuel line within the fuselage must be designed and 
installed to allow a reasonable degree of deformation and stretching 
without leakage.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-15, 
32 FR 13266, Sept. 20, 1967]



Sec. 25.994  Fuel system components.

    Fuel system components in an engine nacelle or in the fuselage must 
be protected from damage that could result in spillage of enough fuel to 
constitute a fire hazard as a result of a wheels-up landing on a paved 
runway under each of the conditions prescribed in Sec. 25.721(b).

[Amdt. No. 25-139, 79 FR 59430, Oct. 2, 2014]



Sec. 25.995  Fuel valves.

    In addition to the requirements of Sec. 25.1189 for shutoff means, 
each fuel valve must--
    (a) [Reserved]
    (b) Be supported so that no loads resulting from their operation or 
from accelerated flight conditions are transmitted to the lines attached 
to the valve.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 
42 FR 15043, Mar. 17, 1977]

[[Page 490]]



Sec. 25.997  Fuel strainer or filter.

    There must be a fuel strainer or filter between the fuel tank outlet 
and the inlet of either the fuel metering device or an engine driven 
positive displacement pump, whichever is nearer the fuel tank outlet. 
This fuel strainer or filter must--
    (a) Be accessible for draining and cleaning and must incorporate a 
screen or element which is easily removable;
    (b) Have a sediment trap and drain except that it need not have a 
drain if the strainer or filter is easily removable for drain purposes;
    (c) Be mounted so that its weight is not supported by the connecting 
lines or by the inlet or outlet connections of the strainer or filter 
itself, unless adequate strength margins under all loading conditions 
are provided in the lines and connections; and
    (d) Have the capacity (with respect to operating limitations 
established for the engine) to ensure that engine fuel system 
functioning is not impaired, with the fuel contaminated to a degree 
(with respect to particle size and density) that is greater than that 
established for the engine in Part 33 of this chapter.

[Amdt. 25-36, 39 FR 35460, Oct. 1, 1974, as amended by Amdt. 25-57, 49 
FR 6848, Feb. 23, 1984]



Sec. 25.999  Fuel system drains.

    (a) Drainage of the fuel system must be accomplished by the use of 
fuel strainer and fuel tank sump drains.
    (b) Each drain required by paragraph (a) of this section must--
    (1) Discharge clear of all parts of the airplane;
    (2) Have manual or automatic means for positive locking in the 
closed position; and
    (3) Have a drain valve--
    (i) That is readily accessible and which can be easily opened and 
closed; and
    (ii) That is either located or protected to prevent fuel spillage in 
the event of a landing with landing gear retracted.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55467, Dec. 20, 1976]



Sec. 25.1001  Fuel jettisoning system.

    (a) A fuel jettisoning system must be installed on each airplane 
unless it is shown that the airplane meets the climb requirements of 
Sec. Sec. 25.119 and 25.121(d) at maximum takeoff weight, less the 
actual or computed weight of fuel necessary for a 15-minute flight 
comprised of a takeoff, go-around, and landing at the airport of 
departure with the airplane configuration, speed, power, and thrust the 
same as that used in meeting the applicable takeoff, approach, and 
landing climb performance requirements of this part.
    (b) If a fuel jettisoning system is required it must be capable of 
jettisoning enough fuel within 15 minutes, starting with the weight 
given in paragraph (a) of this section, to enable the airplane to meet 
the climb requirements of Sec. Sec. 25.119 and 25.121(d), assuming that 
the fuel is jettisoned under the conditions, except weight, found least 
favorable during the flight tests prescribed in paragraph (c) of this 
section.
    (c) Fuel jettisoning must be demonstrated beginning at maximum 
takeoff weight with flaps and landing gear up and in--
    (1) A power-off glide at 1.3 VSR1;
    (2) A climb at the one-engine inoperative best rate-of-climb speed, 
with the critical engine inoperative and the remaining engines at 
maximum continuous power; and
    (3) Level flight at 1.3 V SR1; if the results of the 
tests in the conditions specified in paragraphs (c)(1) and (2) of this 
section show that this condition could be critical.
    (d) During the flight tests prescribed in paragraph (c) of this 
section, it must be shown that--
    (1) The fuel jettisoning system and its operation are free from fire 
hazard;
    (2) The fuel discharges clear of any part of the airplane;
    (3) Fuel or fumes do not enter any parts of the airplane; and
    (4) The jettisoning operation does not adversely affect the 
controllability of the airplane.
    (e) For reciprocating engine powered airplanes, means must be 
provided to prevent jettisoning the fuel in the tanks used for takeoff 
and landing below the level allowing 45 minutes

[[Page 491]]

flight at 75 percent maximum continuous power. However, if there is an 
auxiliary control independent of the main jettisoning control, the 
system may be designed to jettison the remaining fuel by means of the 
auxiliary jettisoning control.
    (f) For turbine engine powered airplanes, means must be provided to 
prevent jettisoning the fuel in the tanks used for takeoff and landing 
below the level allowing climb from sea level to 10,000 feet and 
thereafter allowing 45 minutes cruise at a speed for maximum range. 
However, if there is an auxiliary control independent of the main 
jettisoning control, the system may be designed to jettison the 
remaining fuel by means of the auxiliary jettisoning control.
    (g) The fuel jettisoning valve must be designed to allow flight 
personnel to close the valve during any part of the jettisoning 
operation.
    (h) Unless it is shown that using any means (including flaps, slots, 
and slats) for changing the airflow across or around the wings does not 
adversely affect fuel jettisoning, there must be a placard, adjacent to 
the jettisoning control, to warn flight crewmembers against jettisoning 
fuel while the means that change the airflow are being used.
    (i) The fuel jettisoning system must be designed so that any 
reasonably probable single malfunction in the system will not result in 
a hazardous condition due to unsymmetrical jettisoning of, or inability 
to jettison, fuel.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-18, 
33 FR 12226, Aug. 30, 1968; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984; 
Amdt. 25-108, 67 FR 70827, Nov. 26, 2002]

                               Oil System



Sec. 25.1011  General.

    (a) Each engine must have an independent oil system that can supply 
it with an appropriate quantity of oil at a temperature not above that 
safe for continuous operation.
    (b) The usable oil capacity may not be less than the product of the 
endurance of the airplane under critical operating conditions and the 
approved maximum allowable oil consumption of the engine under the same 
conditions, plus a suitable margin to ensure system circulation. Instead 
of a rational analysis of airplane range for the purpose of computing 
oil requirements for reciprocating engine powered airplanes, the 
following fuel/oil ratios may be used:
    (1) For airplanes without a reserve oil or oil transfer system, a 
fuel/oil ratio of 30:1 by volume.
    (2) For airplanes with either a reserve oil or oil transfer system, 
a fuel/oil ratio of 40:1 by volume.
    (c) Fuel/oil ratios higher than those prescribed in paragraphs 
(b)(1) and (2) of this section may be used if substantiated by data on 
actual engine oil consumption.



Sec. 25.1013  Oil tanks.

    (a) Installation. Each oil tank installation must meet the 
requirements of Sec. 25.967.
    (b) Expansion space. Oil tank expansion space must be provided as 
follows:
    (1) Each oil tank used with a reciprocating engine must have an 
expansion space of not less than the greater of 10 percent of the tank 
capacity or 0.5 gallon, and each oil tank used with a turbine engine 
must have an expansion space of not less than 10 percent of the tank 
capacity.
    (2) Each reserve oil tank not directly connected to any engine may 
have an expansion space of not less than two percent of the tank 
capacity.
    (3) It must be impossible to fill the expansion space inadvertently 
with the airplane in the normal ground attitude.
    (c) Filler connection. Each recessed oil tank filler connection that 
can retain any appreciable quantity of oil must have a drain that 
discharges clear of each part of the airplane. In addition, each oil 
tank filler cap must provide an oil-tight seal.
    (d) Vent. Oil tanks must be vented as follows:
    (1) Each oil tank must be vented from the top part of the expansion 
space so that venting is effective under any normal flight condition.
    (2) Oil tank vents must be arranged so that condensed water vapor 
that might freeze and obstruct the line cannot accumulate at any point.
    (e) Outlet. There must be means to prevent entrance into the tank 
itself,

[[Page 492]]

or into the tank outlet, of any object that might obstruct the flow of 
oil through the system. No oil tank outlet may be enclosed by any screen 
or guard that would reduce the flow of oil below a safe value at any 
operating temperature. There must be a shutoff valve at the outlet of 
each oil tank used with a turbine engine, unless the external portion of 
the oil system (including the oil tank supports) is fireproof.
    (f) Flexible oil tank liners. Each flexible oil tank liner must be 
approved or must be shown to be suitable for the particular application.

[Doc. No. 5066, 29 FR 18291, Dec. 24, as amended by Amdt. 25-19, 33 FR 
15410, Oct. 17, 1968; Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 25-
36, 39 FR 35460, Oct. 1, 1974; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984; 
Amdt. 25-72, 55 FR 29785, July 20, 1990]



Sec. 25.1015  Oil tank tests.

    Each oil tank must be designed and installed so that--
    (a) It can withstand, without failure, each vibration, inertia, and 
fluid load that it may be subjected to in operation; and
    (b) It meets the provisions of Sec. 25.965, except--
    (1) The test pressure--
    (i) For pressurized tanks used with a turbine engine, may not be 
less than 5 p.s.i. plus the maximum operating pressure of the tank 
instead of the pressure specified in Sec. 25.965(a); and
    (ii) For all other tanks may not be less than 5 p.s.i. instead of 
the pressure specified in Sec. 25.965(a); and
    (2) The test fluid must be oil at 250 [deg]F. instead of the fluid 
specified in Sec. 25.965(c).

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-36, 
39 FR 35461, Oct. 1, 1974]



Sec. 25.1017  Oil lines and fittings.

    (a) Each oil line must meet the requirements of Sec. 25.993 and 
each oil line and fitting in any designated fire zone must meet the 
requirements of Sec. 25.1183.
    (b) Breather lines must be arranged so that--
    (1) Condensed water vapor that might freeze and obstruct the line 
cannot accumulate at any point;
    (2) The breather discharge does not constitute a fire hazard if 
foaming occurs or causes emitted oil to strike the pilot's windshield; 
and
    (3) The breather does not discharge into the engine air induction 
system.



Sec. 25.1019  Oil strainer or filter.

    (a) Each turbine engine installation must incorporate an oil 
strainer or filter through which all of the engine oil flows and which 
meets the following requirements:
    (1) Each oil strainer or filter that has a bypass must be 
constructed and installed so that oil will flow at the normal rate 
through the rest of the system with the strainer or filter completely 
blocked.
    (2) The oil strainer or filter must have the capacity (with respect 
to operating limitations established for the engine) to ensure that 
engine oil system functioning is not impaired when the oil is 
contaminated to a degree (with respect to particle size and density) 
that is greater than that established for the engine under Part 33 of 
this chapter.
    (3) The oil strainer or filter, unless it is installed at an oil 
tank outlet, must incorporate an indicator that will indicate 
contamination before it reaches the capacity established in accordance 
with paragraph (a)(2) of this section.
    (4) The bypass of a strainer or filter must be constructed and 
installed so that the release of collected contaminants is minimized by 
appropriate location of the bypass to ensure that collected contaminants 
are not in the bypass flow path.
    (5) An oil strainer or filter that has no bypass, except one that is 
installed at an oil tank outlet, must have a means to connect it to the 
warning system required in Sec. 25.1305(c)(7).
    (b) Each oil strainer or filter in a powerplant installation using 
reciprocating engines must be constructed and installed so that oil will 
flow at the normal rate through the rest of the system with the strainer 
or filter element completely blocked.

[Amdt. 25-36, 39 FR 35461, Oct. 1, 1974, as amended by Amdt. 25-57, 49 
FR 6848, Feb. 23, 1984]

[[Page 493]]



Sec. 25.1021  Oil system drains.

    A drain (or drains) must be provided to allow safe drainage of the 
oil system. Each drain must--
    (a) Be accessible; and
    (b) Have manual or automatic means for positive locking in the 
closed position.

[Amdt. 25-57, 49 FR 6848, Feb. 23, 1984]



Sec. 25.1023  Oil radiators.

    (a) Each oil radiator must be able to withstand, without failure, 
any vibration, inertia, and oil pressure load to which it would be 
subjected in operation.
    (b) Each oil radiator air duct must be located so that, in case of 
fire, flames coming from normal openings of the engine nacelle cannot 
impinge directly upon the radiator.



Sec. 25.1025  Oil valves.

    (a) Each oil shutoff must meet the requirements of Sec. 25.1189.
    (b) The closing of oil shutoff means may not prevent propeller 
feathering.
    (c) Each oil valve must have positive stops or suitable index 
provisions in the ``on'' and ``off'' positions and must be supported so 
that no loads resulting from its operation or from accelerated flight 
conditions are transmitted to the lines attached to the valve.



Sec. 25.1027  Propeller feathering system.

    (a) If the propeller feathering system depends on engine oil, there 
must be means to trap an amount of oil in the tank if the supply becomes 
depleted due to failure of any part of the lubricating system other than 
the tank itself.
    (b) The amount of trapped oil must be enough to accomplish the 
feathering operation and must be available only to the feathering pump.
    (c) The ability of the system to accomplish feathering with the 
trapped oil must be shown. This may be done on the ground using an 
auxiliary source of oil for lubricating the engine during operation.
    (d) Provision must be made to prevent sludge or other foreign matter 
from affecting the safe operation of the propeller feathering system.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55467, Dec. 20, 1976]

                                 Cooling



Sec. 25.1041  General.

    The powerplant and auxiliary power unit cooling provisions must be 
able to maintain the temperatures of powerplant components, engine 
fluids, and auxiliary power unit components and fluids within the 
temperature limits established for these components and fluids, under 
ground, water, and flight operating conditions, and after normal engine 
or auxiliary power unit shutdown, or both.

[Amdt. 25-38, 41 FR 55467, Dec. 20, 1976]



Sec. 25.1043  Cooling tests.

    (a) General. Compliance with Sec. 25.1041 must be shown by tests, 
under critical ground, water, and flight operating conditions. For these 
tests, the following apply:
    (1) If the tests are conducted under conditions deviating from the 
maximum ambient atmospheric temperature, the recorded powerplant 
temperatures must be corrected under paragraphs (c) and (d) of this 
section.
    (2) No corrected temperatures determined under paragraph (a)(1) of 
this section may exceed established limits.
    (3) For reciprocating engines, the fuel used during the cooling 
tests must be the minimum grade approved for the engines, and the 
mixture settings must be those normally used in the flight stages for 
which the cooling tests are conducted. The test procedures must be as 
prescribed in Sec. 25.1045.
    (b) Maximum ambient atmospheric temperature. A maximum ambient 
atmospheric temperature corresponding to sea level conditions of at 
least 100 degrees F must be established. The assumed temperature lapse 
rate is 3.6 degrees F per thousand feet of altitude above sea level 
until a temperature of -69.7 degrees F is reached, above which altitude 
the temperature is considered constant at -69.7 degrees F. However,

[[Page 494]]

for winterization installations, the applicant may select a maximum 
ambient atmospheric temperature corresponding to sea level conditions of 
less than 100 degrees F.
    (c) Correction factor (except cylinder barrels). Unless a more 
rational correction applies, temperatures of engine fluids and 
powerplant components (except cylinder barrels) for which temperature 
limits are established, must be corrected by adding to them the 
difference between the maximum ambient atmospheric temperature and the 
temperature of the ambient air at the time of the first occurrence of 
the maximum component or fluid temperature recorded during the cooling 
test.
    (d) Correction factor for cylinder barrel temperatures. Unless a 
more rational correction applies, cylinder barrel temperatures must be 
corrected by adding to them 0.7 times the difference between the maximum 
ambient atmospheric temperature and the temperature of the ambient air 
at the time of the first occurrence of the maximum cylinder barrel 
temperature recorded during the cooling test.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 
43 FR 2323, Jan. 16, 1978]



Sec. 25.1045  Cooling test procedures.

    (a) Compliance with Sec. 25.1041 must be shown for the takeoff, 
climb, en route, and landing stages of flight that correspond to the 
applicable performance requirements. The cooling tests must be conducted 
with the airplane in the configuration, and operating under the 
conditions, that are critical relative to cooling during each stage of 
flight. For the cooling tests, a temperature is ``stabilized'' when its 
rate of change is less than two degrees F. per minute.
    (b) Temperatures must be stabilized under the conditions from which 
entry is made into each stage of flight being investigated, unless the 
entry condition normally is not one during which component and the 
engine fluid temperatures would stabilize (in which case, operation 
through the full entry condition must be conducted before entry into the 
stage of flight being investigated in order to allow temperatures to 
reach their natural levels at the time of entry). The takeoff cooling 
test must be preceded by a period during which the powerplant component 
and engine fluid temperatures are stabilized with the engines at ground 
idle.
    (c) Cooling tests for each stage of flight must be continued until--
    (1) The component and engine fluid temperatures stabilize;
    (2) The stage of flight is completed; or
    (3) An operating limitation is reached.
    (d) For reciprocating engine powered airplanes, it may be assumed, 
for cooling test purposes, that the takeoff stage of flight is complete 
when the airplane reaches an altitude of 1,500 feet above the takeoff 
surface or reaches a point in the takeoff where the transition from the 
takeoff to the en route configuration is completed and a speed is 
reached at which compliance with Sec. 25.121(c) is shown, whichever 
point is at a higher altitude. The airplane must be in the following 
configuration:
    (1) Landing gear retracted.
    (2) Wing flaps in the most favorable position.
    (3) Cowl flaps (or other means of controlling the engine cooling 
supply) in the position that provides adequate cooling in the hot-day 
condition.
    (4) Critical engine inoperative and its propeller stopped.
    (5) Remaining engines at the maximum continuous power available for 
the altitude.
    (e) For hull seaplanes and amphibians, cooling must be shown during 
taxiing downwind for 10 minutes, at five knots above step speed.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 
49 FR 6848, Feb. 23, 1984]

                            Induction System



Sec. 25.1091  Air induction.

    (a) The air induction system for each engine and auxiliary power 
unit must supply--
    (1) The air required by that engine and auxiliary power unit under 
each operating condition for which certification is requested; and
    (2) The air for proper fuel metering and mixture distribution with 
the induction system valves in any position.

[[Page 495]]

    (b) Each reciprocating engine must have an alternate air source that 
prevents the entry of rain, ice, or any other foreign matter.
    (c) Air intakes may not open within the cowling, unless--
    (1) That part of the cowling is isolated from the engine accessory 
section by means of a fireproof diaphragm; or
    (2) For reciprocating engines, there are means to prevent the 
emergence of backfire flames.
    (d) For turbine engine powered airplanes and airplanes incorporating 
auxiliary power units--
    (1) There must be means to prevent hazardous quantities of fuel 
leakage or overflow from drains, vents, or other components of flammable 
fluid systems from entering the engine or auxiliary power unit intake 
system; and
    (2) The airplane must be designed to prevent water or slush on the 
runway, taxiway, or other airport operating surfaces from being directed 
into the engine or auxiliary power unit air inlet ducts in hazardous 
quantities, and the air inlet ducts must be located or protected so as 
to minimize the ingestion of foreign matter during takeoff, landing, and 
taxiing.
    (e) If the engine induction system contains parts or components that 
could be damaged by foreign objects entering the air inlet, it must be 
shown by tests or, if appropriate, by analysis that the induction system 
design can withstand the foreign object ingestion test conditions of 
Sec. Sec. 33.76, 33.77 and 33.78(a)(1) of this chapter without failure 
of parts or components that could create a hazard.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55467, Dec. 20, 1976; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977; 
Amdt. 25-57, 49 FR 6849, Feb. 23, 1984; Amdt. 25-100, 65 FR 55854, Sept. 
14, 2000]



Sec. 25.1093  Induction system icing protection.

    (a) Reciprocating engines. Each reciprocating engine air induction 
system must have means to prevent and eliminate icing. Unless this is 
done by other means, it must be shown that, in air free of visible 
moisture at a temperature of 30 F., each airplane with altitude engines 
using--
    (1) Conventional venturi carburetors have a preheater that can 
provide a heat rise of 120 F. with the engine at 60 percent of maximum 
continuous power; or
    (2) Carburetors tending to reduce the probability of ice formation 
has a preheater that can provide a heat rise of 100 [deg]F. with the 
engine at 60 percent of maximum continuous power.
    (b) Turbine engines. (1) Each turbine engine must operate throughout 
the flight power range of the engine (including idling), without the 
accumulation of ice on the engine, inlet system components, or airframe 
components that would adversely affect engine operation or cause a 
serious loss of power or thrust--
    (i) Under the icing conditions specified in appendix C, and
    (ii) In falling and blowing snow within the limitations established 
for the airplane for such operation.
    (2) Each turbine engine must idle for 30 minutes on the ground, with 
the air bleed available for engine icing protection at its critical 
condition, without adverse effect, in an atmosphere that is at a 
temperature between 15[deg] and 30 [deg]F (between -9[deg] and -1 
[deg]C) and has a liquid water content not less than 0.3 grams per cubic 
meter in the form of drops having a mean effective diameter not less 
than 20 microns, followed by momentary operation at takeoff power or 
thrust. During the 30 minutes of idle operation, the engine may be run 
up periodically to a moderate power or thrust setting in a manner 
acceptable to the Administrator.
    (c) Supercharged reciprocating engines. For each engine having a 
supercharger to pressurize the air before it enters the carburetor, the 
heat rise in the air caused by that supercharging at any altitude may be 
utilized in determining compliance with paragraph (a) of this section if 
the heat rise utilized is that which will be available, automatically, 
for the applicable altitude and operating condition because of 
supercharging.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55467, Dec. 20, 1976; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977; 
Amdt. 25-57, 49 FR 6849, Feb. 23, 1984; Amdt. 25-72, 55 FR 29785, July 
20, 1990]

[[Page 496]]


    Effective Date Note: At 79 FR 65526, Nov. 4, 2014, Sec. 25.1093 was 
amended by revising paragraph (b), effective Jan. 5, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.1093  Induction system icing protection.

                                * * * * *

    (b) Turbine engines. Except as provided in paragraph (b)(3) of this 
section, each engine, with all icing protection systems operating, must:
    (1) Operate throughout its flight power range, including the minimum 
descent idling speeds, in the icing conditions defined in Appendices C 
and O of this part, and Appendix D of part 33 of this chapter, and in 
falling and blowing snow within the limitations established for the 
airplane for such operation, without the accumulation of ice on the 
engine, inlet system components, or airframe components that would do 
any of the following:
    (i) Adversely affect installed engine operation or cause a sustained 
loss of power or thrust; or an unacceptable increase in gas path 
operating temperature; or an airframe/engine incompatibility; or
    (ii) Result in unacceptable temporary power loss or engine damage; 
or
    (iii) Cause a stall, surge, or flameout or loss of engine 
controllability (for example, rollback).
    (2) Operate at ground idle speed for a minimum of 30 minutes on the 
ground in the following icing conditions shown in Table 1 of this 
section, unless replaced by similar test conditions that are more 
critical. These conditions must be demonstrated with the available air 
bleed for icing protection at its critical condition, without adverse 
effect, followed by an acceleration to takeoff power or thrust in 
accordance with the procedures defined in the airplane flight manual. 
During the idle operation, the engine may be run up periodically to a 
moderate power or thrust setting in a manner acceptable to the 
Administrator. Analysis may be used to show ambient temperatures below 
the tested temperature are less critical. The applicant must document 
the engine run-up procedure (including the maximum time interval between 
run-ups from idle, run-up power setting, and duration at power), the 
associated minimum ambient temperature, and the maximum time interval. 
These conditions must be used in the analysis that establishes the 
airplane operating limitations in accordance with Sec. 25.1521.
    (3) For the purposes of this section, the icing conditions defined 
in appendix O of this part, including the conditions specified in 
Condition 3 of Table 1 of this section, are not applicable to airplanes 
with a maximum takeoff weight equal to or greater than 60,000 pounds.

                                                       Table 1--Icing Conditions for Ground Tests
--------------------------------------------------------------------------------------------------------------------------------------------------------
                                                                            Water concentration       Mean effective particle
              Condition                      Total air temperature               (minimum)                    diameter                 Demonstration
--------------------------------------------------------------------------------------------------------------------------------------------------------
1. Rime ice condition...............  0 to 15 [deg]F (18 to -9 [deg]C)..  Liquid--0.3 g/m\3\.....  15-25 microns................  By test, analysis or
                                                                                                                                   combination of the
                                                                                                                                   two.
2. Glaze ice condition..............  20 to 30 [deg]F (-7 to -1 [deg]C).  Liquid--0.3 g/m\3\.....  15-25 microns................  By test, analysis or
                                                                                                                                   combination of the
                                                                                                                                   two.
3. Large drop condition.............  15 to 30 [deg]F (-9 to -1 [deg]C).  Liquid--0.3 g/m\3\.....  100 microns (minimum)........  By test, analysis or
                                                                                                                                   combination of the
                                                                                                                                   two.
--------------------------------------------------------------------------------------------------------------------------------------------------------

                                * * * * *



Sec. 25.1101  Carburetor air preheater design.

    Each carburetor air preheater must be designed and constructed to--
    (a) Ensure ventilation of the preheater when the engine is operated 
in cold air;
    (b) Allow inspection of the exhaust manifold parts that it 
surrounds; and
    (c) Allow inspection of critical parts of the preheater itself.



Sec. 25.1103  Induction system ducts and air duct systems.

    (a) Each induction system duct upstream of the first stage of the 
engine supercharger and of the auxiliary power unit compressor must have 
a drain to prevent the hazardous accumulation of fuel and moisture in 
the ground attitude. No drain may discharge where it might cause a fire 
hazard.
    (b) Each induction system duct must be--
    (1) Strong enough to prevent induction system failures resulting 
from normal backfire conditions; and

[[Page 497]]

    (2) Fire-resistant if it is in any fire zone for which a fire-
extinguishing system is required, except that ducts for auxiliary power 
units must be fireproof within the auxiliary power unit fire zone.
    (c) Each duct connected to components between which relative motion 
could exist must have means for flexibility.
    (d) For turbine engine and auxiliary power unit bleed air duct 
systems, no hazard may result if a duct failure occurs at any point 
between the air duct source and the airplane unit served by the air.
    (e) Each auxiliary power unit induction system duct must be 
fireproof for a sufficient distance upstream of the auxiliary power unit 
compartment to prevent hot gas reverse flow from burning through 
auxiliary power unit ducts and entering any other compartment or area of 
the airplane in which a hazard would be created resulting from the entry 
of hot gases. The materials used to form the remainder of the induction 
system duct and plenum chamber of the auxiliary power unit must be 
capable of resisting the maximum heat conditions likely to occur.
    (f) Each auxiliary power unit induction system duct must be 
constructed of materials that will not absorb or trap hazardous 
quantities of flammable fluids that could be ignited in the event of a 
surge or reverse flow condition.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 
43 FR 50597, Oct. 30, 1978]



Sec. 25.1105  Induction system screens.

    If induction system screens are used--
    (a) Each screen must be upstream of the carburetor;
    (b) No screen may be in any part of the induction system that is the 
only passage through which air can reach the engine, unless it can be 
deiced by heated air;
    (c) No screen may be deiced by alcohol alone; and
    (d) It must be impossible for fuel to strike any screen.



Sec. 25.1107  Inter-coolers and after-coolers.

    Each inter-cooler and after-cooler must be able to withstand any 
vibration, inertia, and air pressure load to which it would be subjected 
in operation.

                             Exhaust System



Sec. 25.1121  General.

    For powerplant and auxiliary power unit installations the following 
apply:
    (a) Each exhaust system must ensure safe disposal of exhaust gases 
without fire hazard or carbon monoxide contamination in any personnel 
compartment. For test purposes, any acceptable carbon monoxide detection 
method may be used to show the absence of carbon monoxide.
    (b) Each exhaust system part with a surface hot enough to ignite 
flammable fluids or vapors must be located or shielded so that leakage 
from any system carrying flammable fluids or vapors will not result in a 
fire caused by impingement of the fluids or vapors on any part of the 
exhaust system including shields for the exhaust system.
    (c) Each component that hot exhaust gases could strike, or that 
could be subjected to high temperatures from exhaust system parts, must 
be fireproof. All exhaust system components must be separated by 
fireproof shields from adjacent parts of the airplane that are outside 
the engine and auxiliary power unit compartments.
    (d) No exhaust gases may discharge so as to cause a fire hazard with 
respect to any flammable fluid vent or drain.
    (e) No exhaust gases may discharge where they will cause a glare 
seriously affecting pilot vision at night.
    (f) Each exhaust system component must be ventilated to prevent 
points of excessively high temperature.
    (g) Each exhaust shroud must be ventilated or insulated to avoid, 
during normal operation, a temperature high enough to ignite any 
flammable fluids or vapors external to the shroud.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 
42 FR 15043, Mar. 17, 1977]

[[Page 498]]



Sec. 25.1123  Exhaust piping.

    For powerplant and auxiliary power unit installations, the following 
apply:
    (a) Exhaust piping must be heat and corrosion resistant, and must 
have provisions to prevent failure due to expansion by operating 
temperatures.
    (b) Piping must be supported to withstand any vibration and inertia 
loads to which it would be subjected in operation; and
    (c) Piping connected to components between which relative motion 
could exist must have means for flexibility.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 
42 FR 15044, Mar. 17, 1977]



Sec. 25.1125  Exhaust heat exchangers.

    For reciprocating engine powered airplanes, the following apply:
    (a) Each exhaust heat exchanger must be constructed and installed to 
withstand each vibration, inertia, and other load to which it would be 
subjected in operation. In addition--
    (1) Each exchanger must be suitable for continued operation at high 
temperatures and resistant to corrosion from exhaust gases;
    (2) There must be means for the inspection of the critical parts of 
each exchanger;
    (3) Each exchanger must have cooling provisions wherever it is 
subject to contact with exhaust gases; and
    (4) No exhaust heat exchanger or muff may have any stagnant areas or 
liquid traps that would increase the probability of ignition of 
flammable fluids or vapors that might be present in case of the failure 
or malfunction of components carrying flammable fluids.
    (b) If an exhaust heat exchanger is used for heating ventilating 
air--
    (1) There must be a secondary heat exchanger between the primary 
exhaust gas heat exchanger and the ventilating air system; or
    (2) Other means must be used to preclude the harmful contamination 
of the ventilating air.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55467, Dec. 20, 1976]



Sec. 25.1127  Exhaust driven turbo-superchargers.

    (a) Each exhaust driven turbo-supercharger must be approved or shown 
to be suitable for the particular application. It must be installed and 
supported to ensure safe operation between normal inspections and 
overhauls. In addition, there must be provisions for expansion and 
flexibility between exhaust conduits and the turbine.
    (b) There must be provisions for lubricating the turbine and for 
cooling turbine parts where temperatures are critical.
    (c) If the normal turbo-supercharger control system malfunctions, 
the turbine speed may not exceed its maximum allowable value. Except for 
the waste gate operating components, the components provided for meeting 
this requirement must be independent of the normal turbo-supercharger 
controls.

                   Powerplant Controls and Accessories



Sec. 25.1141  Powerplant controls: general.

    Each powerplant control must be located, arranged, and designed 
under Sec. Sec. 25.777 through 25.781 and marked under Sec. 25.1555. 
In addition, it must meet the following requirements:
    (a) Each control must be located so that it cannot be inadvertently 
operated by persons entering, leaving, or moving normally in, the 
cockpit.
    (b) Each flexible control must be approved or must be shown to be 
suitable for the particular application.
    (c) Each control must have sufficient strength and rigidity to 
withstand operating loads without failure and without excessive 
deflection.
    (d) Each control must be able to maintain any set position without 
constant attention by flight crewmembers and without creep due to 
control loads or vibration.
    (e) The portion of each powerplant control located in a designated 
fire zone that is required to be operated in the event of fire must be 
at least fire resistant.
    (f) For powerplant valve controls located in the flight deck there 
must be a means:

[[Page 499]]

    (1) For the flightcrew to select each intended position or function 
of the valve; and
    (2) To indicate to the flightcrew:
    (i) The selected position or function of the valve; and
    (ii) When the valve has not responded as intended to the selected 
position or function.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 
42 FR 15044, Mar. 17, 1977; Amdt. 25-72, 55 FR 29785, July 20, 1990; 
Amdt. 25-115, 69 FR 40527, July 2, 2004]



Sec. 25.1142  Auxiliary power unit controls.

    Means must be provided on the flight deck for starting, stopping, 
and emergency shutdown of each installed auxiliary power unit.

[Amdt. 25-46, 43 FR 50598, Oct. 30, 1978]



Sec. 25.1143  Engine controls.

    (a) There must be a separate power or thrust control for each 
engine.
    (b) Power and thrust controls must be arranged to allow--
    (1) Separate control of each engine; and
    (2) Simultaneous control of all engines.
    (c) Each power and thrust control must provide a positive and 
immediately responsive means of controlling its engine.
    (d) For each fluid injection (other than fuel) system and its 
controls not provided and approved as part of the engine, the applicant 
must show that the flow of the injection fluid is adequately controlled.
    (e) If a power or thrust control incorporates a fuel shutoff 
feature, the control must have a means to prevent the inadvertent 
movement of the control into the shutoff position. The means must--
    (1) Have a positive lock or stop at the idle position; and
    (2) Require a separate and distinct operation to place the control 
in the shutoff position.

[Amdt. 25-23, 35 FR 5677, Apr. 8, 1970, as amended by Amdt. 25-38, 41 FR 
55467, Dec. 20, 1976; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984]



Sec. 25.1145  Ignition switches.

    (a) Ignition switches must control each engine ignition circuit on 
each engine.
    (b) There must be means to quickly shut off all ignition by the 
grouping of switches or by a master ignition control.
    (c) Each group of ignition switches, except ignition switches for 
turbine engines for which continuous ignition is not required, and each 
master ignition control must have a means to prevent its inadvertent 
operation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 
42 FR 15044 Mar. 17, 1977]



Sec. 25.1147  Mixture controls.

    (a) If there are mixture controls, each engine must have a separate 
control. The controls must be grouped and arranged to allow--
    (1) Separate control of each engine; and
    (2) Simultaneous control of all engines.
    (b) Each intermediate position of the mixture controls that 
corresponds to a normal operating setting must be identifiable by feel 
and sight.
    (c) The mixture controls must be accessible to both pilots. However, 
if there is a separate flight engineer station with a control panel, the 
controls need be accessible only to the flight engineer.



Sec. 25.1149  Propeller speed and pitch controls.

    (a) There must be a separate propeller speed and pitch control for 
each propeller.
    (b) The controls must be grouped and arranged to allow--
    (1) Separate control of each propeller; and
    (2) Simultaneous control of all propellers.
    (c) The controls must allow synchronization of all propellers.
    (d) The propeller speed and pitch controls must be to the right of, 
and at least one inch below, the pilot's throttle controls.

[[Page 500]]



Sec. 25.1153  Propeller feathering controls.

    (a) There must be a separate propeller feathering control for each 
propeller. The control must have means to prevent its inadvertent 
operation.
    (b) If feathering is accomplished by movement of the propeller pitch 
or speed control lever, there must be means to prevent the inadvertent 
movement of this lever to the feathering position during normal 
operation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 
32 FR 6913, May 5, 1967]



Sec. 25.1155  Reverse thrust and propeller pitch settings below the
flight regime.

    Each control for reverse thrust and for propeller pitch settings 
below the flight regime must have means to prevent its inadvertent 
operation. The means must have a positive lock or stop at the flight 
idle position and must require a separate and distinct operation by the 
crew to displace the control from the flight regime (forward thrust 
regime for turbojet powered airplanes).

[Amdt. 25-11, 32 FR 6913, May 5, 1967]



Sec. 25.1157  Carburetor air temperature controls.

    There must be a separate carburetor air temperature control for each 
engine.



Sec. 25.1159  Supercharger controls.

    Each supercharger control must be accessible to the pilots or, if 
there is a separate flight engineer station with a control panel, to the 
flight engineer.



Sec. 25.1161  Fuel jettisoning system controls.

    Each fuel jettisoning system control must have guards to prevent 
inadvertent operation. No control may be near any fire extinguisher 
control or other control used to combat fire.



Sec. 25.1163  Powerplant accessories.

    (a) Each engine mounted accessory must--
    (1) Be approved for mounting on the engine involved;
    (2) Use the provisions on the engine for mounting; and
    (3) Be sealed to prevent contamination of the engine oil system and 
the accessory system.
    (b) Electrical equipment subject to arcing or sparking must be 
installed to minimize the probability of contact with any flammable 
fluids or vapors that might be present in a free state.
    (c) If continued rotation of an engine-driven cabin supercharger or 
of any remote accessory driven by the engine is hazardous if 
malfunctioning occurs, there must be means to prevent rotation without 
interfering with the continued operation of the engine.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 
49 FR 6849, Feb. 23, 1984]



Sec. 25.1165  Engine ignition systems.

    (a) Each battery ignition system must be supplemented by a generator 
that is automatically available as an alternate source of electrical 
energy to allow continued engine operation if any battery becomes 
depleted.
    (b) The capacity of batteries and generators must be large enough to 
meet the simultaneous demands of the engine ignition system and the 
greatest demands of any electrical system components that draw 
electrical energy from the same source.
    (c) The design of the engine ignition system must account for--
    (1) The condition of an inoperative generator;
    (2) The condition of a completely depleted battery with the 
generator running at its normal operating speed; and
    (3) The condition of a completely depleted battery with the 
generator operating at idling speed, if there is only one battery.
    (d) Magneto ground wiring (for separate ignition circuits) that lies 
on the engine side of the fire wall, must be installed, located, or 
protected, to minimize the probability of simultaneous failure of two or 
more wires as a result of mechanical damage, electrical faults, or other 
cause.
    (e) No ground wire for any engine may be routed through a fire zone 
of another engine unless each part of that wire within that zone is 
fireproof.

[[Page 501]]

    (f) Each ignition system must be independent of any electrical 
circuit, not used for assisting, controlling, or analyzing the operation 
of that system.
    (g) There must be means to warn appropriate flight crewmembers if 
the malfunctioning of any part of the electrical system is causing the 
continuous discharge of any battery necessary for engine ignition.
    (h) Each engine ignition system of a turbine powered airplane must 
be considered an essential electrical load.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5677, Apr. 8, 1970; Amdt. 25-72, 55 FR 29785, July 20, 1990]



Sec. 25.1167  Accessory gearboxes.

    For airplanes equipped with an accessory gearbox that is not 
certificated as part of an engine--
    (a) The engine with gearbox and connecting transmissions and shafts 
attached must be subjected to the tests specified in Sec. 33.49 or 
Sec. 33.87 of this chapter, as applicable;
    (b) The accessory gearbox must meet the requirements of Sec. Sec. 
33.25 and 33.53 or 33.91 of this chapter, as applicable; and
    (c) Possible misalignments and torsional loadings of the gearbox, 
transmission, and shaft system, expected to result under normal 
operating conditions must be evaluated.

[Amdt. 25-38, 41 FR 55467, Dec. 20, 1976]

                       Powerplant Fire Protection



Sec. 25.1181  Designated fire zones; regions included.

    (a) Designated fire zones are--
    (1) The engine power section;
    (2) The engine accessory section;
    (3) Except for reciprocating engines, any complete powerplant 
compartment in which no isolation is provided between the engine power 
section and the engine accessory section;
    (4) Any auxiliary power unit compartment;
    (5) Any fuel-burning heater and other combustion equipment 
installation described in Sec. 25.859;
    (6) The compressor and accessory sections of turbine engines; and
    (7) Combustor, turbine, and tailpipe sections of turbine engine 
installations that contain lines or components carrying flammable fluids 
or gases.
    (b) Each designated fire zone must meet the requirements of 
Sec. Sec. 25.863, 25.865, 25.867, 25.869, and 25.1185 through 25.1203.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 
32 FR 6913, May 5, 1967; Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 
25-72, 55 FR 29785, July 20, 1990; Amdt. 25-115, 69 FR 40527, July 2, 
2004]



Sec. 25.1182  Nacelle areas behind firewalls, and engine pod attaching
structures containing flammable fluid lines.

    (a) Each nacelle area immediately behind the firewall, and each 
portion of any engine pod attaching structure containing flammable fluid 
lines, must meet each requirement of Sec. Sec. 25.1103(b), 25.1165 (d) 
and (e), 25.1183, 25.1185(c), 25.1187, 25.1189, and 25.1195 through 
25.1203, including those concerning designated fire zones. However, 
engine pod attaching structures need not contain fire detection or 
extinguishing means.
    (b) For each area covered by paragraph (a) of this section that 
contains a retractable landing gear, compliance with that paragraph need 
only be shown with the landing gear retracted.

[Amdt. 25-11, 32 FR 6913, May 5, 1967]



Sec. 25.1183  Flammable fluid-carrying components.

    (a) Except as provided in paragraph (b) of this section, each line, 
fitting, and other component carrying flammable fluid in any area 
subject to engine fire conditions, and each component which conveys or 
contains flammable fluid in a designated fire zone must be fire 
resistant, except that flammable fluid tanks and supports in a 
designated fire zone must be fireproof or be enclosed by a fireproof 
shield unless damage by fire to any non-fireproof part will not cause 
leakage or spillage of flammable fluid. Components must be shielded or 
located to safeguard against the ignition of leaking flammable fluid. An 
integral oil sump of less than 25-quart capacity on a reciprocating 
engine need not be fireproof nor be enclosed by a fireproof shield.
    (b) Paragraph (a) of this section does not apply to--

[[Page 502]]

    (1) Lines, fittings, and components which are already approved as 
part of a type certificated engine; and
    (2) Vent and drain lines, and their fittings, whose failure will not 
result in, or add to, a fire hazard.
    (c) All components, including ducts, within a designated fire zone 
must be fireproof if, when exposed to or damaged by fire, they could--
    (1) Result in fire spreading to other regions of the airplane; or
    (2) Cause unintentional operation of, or inability to operate, 
essential services or equipment.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 
32 FR 6913, May 5, 1967; Amdt. 25-36, 39 FR 35461, Oct. 1, 1974; Amdt. 
25-57, 49 FR 6849, Feb. 23, 1984; Amdt. 25-101, 65 FR 79710, Dec. 19, 
2000]



Sec. 25.1185  Flammable fluids.

    (a) Except for the integral oil sumps specified in Sec. 25.1183(a), 
no tank or reservoir that is a part of a system containing flammable 
fluids or gases may be in a designated fire zone unless the fluid 
contained, the design of the system, the materials used in the tank, the 
shut-off means, and all connections, lines, and control provide a degree 
of safety equal to that which would exist if the tank or reservoir were 
outside such a zone.
    (b) There must be at least one-half inch of clear airspace between 
each tank or reservoir and each firewall or shroud isolating a 
designated fire zone.
    (c) Absorbent materials close to flammable fluid system components 
that might leak must be covered or treated to prevent the absorption of 
hazardous quantities of fluids.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964 as amended by Amdt. 25-19, 33 
FR 15410, Oct. 17, 1968; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998]



Sec. 25.1187  Drainage and ventilation of fire zones.

    (a) There must be complete drainage of each part of each designated 
fire zone to minimize the hazards resulting from failure or 
malfunctioning of any component containing flammable fluids. The 
drainage means must be--
    (1) Effective under conditions expected to prevail when drainage is 
needed; and
    (2) Arranged so that no discharged fluid will cause an additional 
fire hazard.
    (b) Each designated fire zone must be ventilated to prevent the 
accumulation of flammable vapors.
    (c) No ventilation opening may be where it would allow the entry of 
flammable fluids, vapors, or flame from other zones.
    (d) Each ventilation means must be arranged so that no discharged 
vapors will cause an additional fire hazard.
    (e) Unless the extinguishing agent capacity and rate of discharge 
are based on maximum air flow through a zone, there must be means to 
allow the crew to shut off sources of forced ventilation to any fire 
zone except the engine power section of the nacelle and the combustion 
heater ventilating air ducts.



Sec. 25.1189  Shutoff means.

    (a) Each engine installation and each fire zone specified in Sec. 
25.1181(a)(4) and (5) must have a means to shut off or otherwise prevent 
hazardous quantities of fuel, oil, deicer, and other flammable fluids, 
from flowing into, within, or through any designated fire zone, except 
that shutoff means are not required for--
    (1) Lines, fittings, and components forming an integral part of an 
engine; and
    (2) Oil systems for turbine engine installations in which all 
components of the system in a designated fire zone, including oil tanks, 
are fireproof or located in areas not subject to engine fire conditions.
    (b) The closing of any fuel shutoff valve for any engine may not 
make fuel unavailable to the remaining engines.
    (c) Operation of any shutoff may not interfere with the later 
emergency operation of other equipment, such as the means for feathering 
the propeller.
    (d) Each flammable fluid shutoff means and control must be fireproof 
or must be located and protected so that any fire in a fire zone will 
not affect its operation.
    (e) No hazardous quantity of flammable fluid may drain into any 
designated fire zone after shutoff.

[[Page 503]]

    (f) There must be means to guard against inadvertent operation of 
the shutoff means and to make it possible for the crew to reopen the 
shutoff means in flight after it has been closed.
    (g) Each tank-to-engine shutoff valve must be located so that the 
operation of the valve will not be affected by powerplant or engine 
mount structural failure.
    (h) Each shutoff valve must have a means to relieve excessive 
pressure accumulation unless a means for pressure relief is otherwise 
provided in the system.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5677, Apr. 8, 1970; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984]



Sec. 25.1191  Firewalls.

    (a) Each engine, auxiliary power unit, fuel-burning heater, other 
combustion equipment intended for operation in flight, and the 
combustion, turbine, and tailpipe sections of turbine engines, must be 
isolated from the rest of the airplane by firewalls, shrouds, or 
equivalent means.
    (b) Each firewall and shroud must be--
    (1) Fireproof;
    (2) Constructed so that no hazardous quantity of air, fluid, or 
flame can pass from the compartment to other parts of the airplane;
    (3) Constructed so that each opening is sealed with close fitting 
fireproof grommets, bushings, or firewall fittings; and
    (4) Protected against corrosion.



Sec. 25.1192  Engine accessory section diaphragm.

    For reciprocating engines, the engine power section and all portions 
of the exhaust system must be isolated from the engine accessory 
compartment by a diaphragm that complies with the firewall requirements 
of Sec. 25.1191.

[Amdt. 25-23, 35 FR 5678, Apr. 8, 1970]



Sec. 25.1193  Cowling and nacelle skin.

    (a) Each cowling must be constructed and supported so that it can 
resist any vibration, inertia, and air load to which it may be subjected 
in operation.
    (b) Cowling must meet the drainage and ventilation requirements of 
Sec. 25.1187.
    (c) On airplanes with a diaphragm isolating the engine power section 
from the engine accessory section, each part of the accessory section 
cowling subject to flame in case of fire in the engine power section of 
the powerplant must--
    (1) Be fireproof; and
    (2) Meet the requirements of Sec. 25.1191.
    (d) Each part of the cowling subject to high temperatures due to its 
nearness to exhaust system parts or exhaust gas impingement must be 
fireproof.
    (e) Each airplane must--
    (1) Be designed and constructed so that no fire originating in any 
fire zone can enter, either through openings or by burning through 
external skin, any other zone or region where it would create additional 
hazards;
    (2) Meet paragraph (e)(1) of this section with the landing gear 
retracted (if applicable); and
    (3) Have fireproof skin in areas subject to flame if a fire starts 
in the engine power or accessory sections.



Sec. 25.1195  Fire extinguishing systems.

    (a) Except for combustor, turbine, and tail pipe sections of turbine 
engine installations that contain lines or components carrying flammable 
fluids or gases for which it is shown that a fire originating in these 
sections can be controlled, there must be a fire extinguisher system 
serving each designated fire zone.
    (b) The fire extinguishing system, the quantity of the extinguishing 
agent, the rate of discharge, and the discharge distribution must be 
adequate to extinguish fires. It must be shown by either actual or 
simulated flights tests that under critical airflow conditions in flight 
the discharge of the extinguishing agent in each designated fire zone 
specified in paragraph (a) of this section will provide an agent 
concentration capable of extinguishing fires in that zone and of 
minimizing the probability of reignition. An individual ``one-shot'' 
system may be used for auxiliary power units, fuel burning heaters, and 
other combustion equipment. For each other designated fire

[[Page 504]]

zone, two discharges must be provided each of which produces adequate 
agent concentration.
    (c) The fire extinguishing system for a nacelle must be able to 
simultaneously protect each zone of the nacelle for which protection is 
provided.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 
43 FR 50598, Oct. 30, 1978]



Sec. 25.1197  Fire extinguishing agents.

    (a) Fire extinguishing agents must--
    (1) Be capable of extinguishing flames emanating from any burning of 
fluids or other combustible materials in the area protected by the fire 
extinguishing system; and
    (2) Have thermal stability over the temperature range likely to be 
experienced in the compartment in which they are stored.
    (b) If any toxic extinguishing agent is used, provisions must be 
made to prevent harmful concentrations of fluid or fluid vapors (from 
leakage during normal operation of the airplane or as a result of 
discharging the fire extinguisher on the ground or in flight) from 
entering any personnel compartment, even though a defect may exist in 
the extinguishing system. This must be shown by test except for built-in 
carbon dioxide fuselage compartment fire extinguishing systems for 
which--
    (1) Five pounds or less of carbon dioxide will be discharged, under 
established fire control procedures, into any fuselage compartment; or
    (2) There is protective breathing equipment for each flight 
crewmember on flight deck duty.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55467, Dec. 20, 1976; Amdt. 25-40, 42 FR 15044, Mar. 17, 1977]



Sec. 25.1199  Extinguishing agent containers.

    (a) Each extinguishing agent container must have a pressure relief 
to prevent bursting of the container by excessive internal pressures.
    (b) The discharge end of each discharge line from a pressure relief 
connection must be located so that discharge of the fire extinguishing 
agent would not damage the airplane. The line must also be located or 
protected to prevent clogging caused by ice or other foreign matter.
    (c) There must be a means for each fire extinguishing agent 
container to indicate that the container has discharged or that the 
charging pressure is below the established minimum necessary for proper 
functioning.
    (d) The temperature of each container must be maintained, under 
intended operating conditions, to prevent the pressure in the container 
from--
    (1) Falling below that necessary to provide an adequate rate of 
discharge; or
    (2) Rising high enough to cause premature discharge.
    (e) If a pyrotechnic capsule is used to discharge the extinguishing 
agent, each container must be installed so that temperature conditions 
will not cause hazardous deterioration of the pyrotechnic capsule.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5678, Apr. 8, 1970; Amdt. 25-40, 42 FR 15044, Mar. 17, 1977]



Sec. 25.1201  Fire extinguishing system materials.

    (a) No material in any fire extinguishing system may react 
chemically with any extinguishing agent so as to create a hazard.
    (b) Each system component in an engine compartment must be 
fireproof.



Sec. 25.1203  Fire detector system.

    (a) There must be approved, quick acting fire or overheat detectors 
in each designated fire zone, and in the combustion, turbine, and 
tailpipe sections of turbine engine installations, in numbers and 
locations ensuring prompt detection of fire in those zones.
    (b) Each fire detector system must be constructed and installed so 
that--
    (1) It will withstand the vibration, inertia, and other loads to 
which it may be subjected in operation;
    (2) There is a means to warn the crew in the event that the sensor 
or associated wiring within a designated fire zone is severed at one 
point, unless the system continues to function as a satisfactory 
detection system after the severing; and
    (3) There is a means to warn the crew in the event of a short 
circuit in the sensor or associated wiring within a

[[Page 505]]

designated fire zone, unless the system continues to function as a 
satisfactory detection system after the short circuit.
    (c) No fire or overheat detector may be affected by any oil, water, 
other fluids or fumes that might be present.
    (d) There must be means to allow the crew to check, in flight, the 
functioning of each fire or overheat detector electric circuit.
    (e) Components of each fire or overheat detector system in a fire 
zone must be fire-resistant.
    (f) No fire or overheat detector system component for any fire zone 
may pass through another fire zone, unless--
    (1) It is protected against the possibility of false warnings 
resulting from fires in zones through which it passes; or
    (2) Each zone involved is simultaneously protected by the same 
detector and extinguishing system.
    (g) Each fire detector system must be constructed so that when it is 
in the configuration for installation it will not exceed the alarm 
activation time approved for the detectors using the response time 
criteria specified in the appropriate Technical Standard Order for the 
detector.
    (h) EWIS for each fire or overheat detector system in a fire zone 
must meet the requirements of Sec. 25.1731.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5678, Apr. 8, 1970; Amdt. 25-26, 36 FR 5493, Mar. 24, 1971; Amdt. 
25-123, 72 FR 63405, Nov. 8, 2007]



Sec. 25.1207  Compliance.

    Unless otherwise specified, compliance with the requirements of 
Sec. Sec. 25.1181 through 25.1203 must be shown by a full scale fire 
test or by one or more of the following methods:
    (a) Tests of similar powerplant configurations;
    (b) Tests of components;
    (c) Service experience of aircraft with similar powerplant 
configurations;
    (d) Analysis.

[Amdt. 25-46, 43 FR 50598, Oct. 30, 1978]



                           Subpart F_Equipment

                                 General



Sec. 25.1301  Function and installation.

    (a) Each item of installed equipment must--
    (1) Be of a kind and design appropriate to its intended function;
    (2) Be labeled as to its identification, function, or operating 
limitations, or any applicable combination of these factors;
    (3) Be installed according to limitations specified for that 
equipment; and
    (4) Function properly when installed.
    (b) EWIS must meet the requirements of subpart H of this part.

[Doc.. No. 5066, Amdt. 1-6, 29 FR 18333, Dec. 24, 1964, as amended by 
Amdt. 25-123, 72 FR 63405, Nov. 8, 2007]



Sec. 25.1302  Installed systems and equipment for use by the flightcrew.

    This section applies to installed systems and equipment intended for 
flightcrew members' use in operating the airplane from their normally 
seated positions on the flight deck. The applicant must show that these 
systems and installed equipment, individually and in combination with 
other such systems and equipment, are designed so that qualified 
flightcrew members trained in their use can safely perform all of the 
tasks associated with the systems' and equipment's intended functions. 
Such installed equipment and systems must meet the following 
requirements:
    (a) Flight deck controls must be installed to allow accomplishment 
of all the tasks required to safely perform the equipment's intended 
function, and information must be provided to the flightcrew that is 
necessary to accomplish the defined tasks.
    (b) Flight deck controls and information intended for the 
flightcrew's use must:
    (1) Be provided in a clear and unambiguous manner at a resolution 
and precision appropriate to the task;
    (2) Be accessible and usable by the flightcrew in a manner 
consistent with the urgency, frequency, and duration of their tasks; and

[[Page 506]]

    (3) Enable flightcrew awareness, if awareness is required for safe 
operation, of the effects on the airplane or systems resulting from 
flightcrew actions.
    (c) Operationally-relevant behavior of the installed equipment must 
be:
    (1) Predictable and unambiguous; and
    (2) Designed to enable the flightcrew to intervene in a manner 
appropriate to the task.
    (d) To the extent practicable, installed equipment must incorporate 
means to enable the flightcrew to manage errors resulting from the kinds 
of flightcrew interactions with the equipment that can be reasonably 
expected in service. This paragraph does not apply to any of the 
following:
    (1) Skill-related errors associated with manual control of the 
airplane;
    (2) Errors that result from decisions, actions, or omissions 
committed with malicious intent;
    (3) Errors arising from a crewmember's reckless decisions, actions, 
or omissions reflecting a substantial disregard for safety; and
    (4) Errors resulting from acts or threats of violence, including 
actions taken under duress.

[Doc. No. FAA-2010-1175, 78 FR 25846, May 3, 2013]



Sec. 25.1303  Flight and navigation instruments.

    (a) The following flight and navigation instruments must be 
installed so that the instrument is visible from each pilot station:
    (1) A free air temperature indicator or an air-temperature indicator 
which provides indications that are convertible to free-air temperature.
    (2) A clock displaying hours, minutes, and seconds with a sweep-
second pointer or digital presentation.
    (3) A direction indicator (nonstabilized magnetic compass).
    (b) The following flight and navigation instruments must be 
installed at each pilot station:
    (1) An airspeed indicator. If airspeed limitations vary with 
altitude, the indicator must have a maximum allowable airspeed indicator 
showing the variation of VMO with altitude.
    (2) An altimeter (sensitive).
    (3) A rate-of-climb indicator (vertical speed).
    (4) A gyroscopic rate-of-turn indicator combined with an integral 
slip-skid indicator (turn-and-bank indicator) except that only a slip-
skid indicator is required on large airplanes with a third attitude 
instrument system useable through flight attitudes of 360[deg] of pitch 
and roll and installed in accordance with Sec. 121.305(k) of this 
title.
    (5) A bank and pitch indicator (gyroscopically stabilized).
    (6) A direction indicator (gyroscopically stabilized, magnetic or 
nonmagnetic).
    (c) The following flight and navigation instruments are required as 
prescribed in this paragraph:
    (1) A speed warning device is required for turbine engine powered 
airplanes and for airplanes with VMO/MMO greater than 0.8 VDF/MDF or 0.8 
V D/MD. The speed warning device must give effective aural warning 
(differing distinctively from aural warnings used for other purposes) to 
the pilots, whenever the speed exceeds VMO plus 6 knots or MMO +0.01. 
The upper limit of the production tolerance for the warning device may 
not exceed the prescribed warning speed.
    (2) A machmeter is required at each pilot station for airplanes with 
compressibility limitations not otherwise indicated to the pilot by the 
airspeed indicating system required under paragraph (b)(1) of this 
section.

[Amdt. 25-23, 35 FR 5678, Apr. 8, 1970, as amended by Amdt. 25-24, 35 FR 
7108, May 6, 1970; Amdt. 25-38, 41 FR 55467, Dec. 20, 1976; Amdt. 25-90, 
62 FR 13253, Mar. 19, 1997]



Sec. 25.1305  Powerplant instruments.

    The following are required powerplant instruments:
    (a) For all airplanes. (1) A fuel pressure warning means for each 
engine, or a master warning means for all engines with provision for 
isolating the individual warning means from the master warning means.
    (2) A fuel quantity indicator for each fuel tank.
    (3) An oil quantity indicator for each oil tank.

[[Page 507]]

    (4) An oil pressure indicator for each independent pressure oil 
system of each engine.
    (5) An oil pressure warning means for each engine, or a master 
warning means for all engines with provision for isolating the 
individual warning means from the master warning means.
    (6) An oil temperature indicator for each engine.
    (7) Fire-warning devices that provide visual and audible warning.
    (8) An augmentation liquid quantity indicator (appropriate for the 
manner in which the liquid is to be used in operation) for each tank.
    (b) For reciprocating engine-powered airplanes. In addition to the 
powerplant instruments required by paragraph (a) of this section, the 
following powerplant instruments are required:
    (1) A carburetor air temperature indicator for each engine.
    (2) A cylinder head temperature indicator for each air-cooled 
engine.
    (3) A manifold pressure indicator for each engine.
    (4) A fuel pressure indicator (to indicate the pressure at which the 
fuel is supplied) for each engine.
    (5) A fuel flowmeter, or fuel mixture indicator, for each engine 
without an automatic altitude mixture control.
    (6) A tachometer for each engine.
    (7) A device that indicates, to the flight crew (during flight), any 
change in the power output, for each engine with--
    (i) An automatic propeller feathering system, whose operation is 
initiated by a power output measuring system; or
    (ii) A total engine piston displacement of 2,000 cubic inches or 
more.
    (8) A means to indicate to the pilot when the propeller is in 
reverse pitch, for each reversing propeller.
    (c) For turbine engine-powered airplanes. In addition to the 
powerplant instruments required by paragraph (a) of this section, the 
following powerplant instruments are required:
    (1) A gas temperature indicator for each engine.
    (2) A fuel flowmeter indicator for each engine.
    (3) A tachometer (to indicate the speed of the rotors with 
established limiting speeds) for each engine.
    (4) A means to indicate, to the flight crew, the operation of each 
engine starter that can be operated continuously but that is neither 
designed for continuous operation nor designed to prevent hazard if it 
failed.
    (5) An indicator to indicate the functioning of the powerplant ice 
protection system for each engine.
    (6) An indicator for the fuel strainer or filter required by Sec. 
25.997 to indicate the occurrence of contamination of the strainer or 
filter before it reaches the capacity established in accordance with 
Sec. 25.997(d).
    (7) A warning means for the oil strainer or filter required by Sec. 
25.1019, if it has no bypass, to warn the pilot of the occurrence of 
contamination of the strainer or filter screen before it reaches the 
capacity established in accordance with Sec. 25.1019(a)(2).
    (8) An indicator to indicate the proper functioning of any heater 
used to prevent ice clogging of fuel system components.
    (d) For turbojet engine powered airplanes. In addition to the 
powerplant instruments required by paragraphs (a) and (c) of this 
section, the following powerplant instruments are required:
    (1) An indicator to indicate thrust, or a parameter that is directly 
related to thrust, to the pilot. The indication must be based on the 
direct measurement of thrust or of parameters that are directly related 
to thrust. The indicator must indicate a change in thrust resulting from 
any engine malfunction, damage, or deterioration.
    (2) A position indicating means to indicate to the flightcrew when 
the thrust reversing device--
    (i) Is not in the selected position, and
    (ii) Is in the reverse thrust position, for each engine using a 
thrust reversing device.
    (3) An indicator to indicate rotor system unbalance.
    (e) For turbopropeller-powered airplanes. In addition to the 
powerplant instruments required by paragraphs (a) and (c) of this 
section, the following powerplant instruments are required:
    (1) A torque indicator for each engine.

[[Page 508]]

    (2) Position indicating means to indicate to the flight crew when 
the propeller blade angle is below the flight low pitch position, for 
each propeller.
    (f) For airplanes equipped with fluid systems (other than fuel) for 
thrust or power augmentation, an approved means must be provided to 
indicate the proper functioning of that system to the flight crew.

[Amdt. 25-23, 35 FR 5678, Apr. 8, 1970, as amended by Amdt. 25-35, 39 FR 
1831, Jan. 15, 1974; Amdt. 25-36, 39 FR 35461, Oct. 1, 1974; Amdt. 25-
38, 41 FR 55467, Dec. 20, 1976; Amdt. 25-54, 45 FR 60173, Sept. 11, 
1980; Amdt. 25-72, 55 FR 29785, July 20, 1990; Amdt. 25-115, 69 FR 
40527, July 2, 2004]



Sec. 25.1307  Miscellaneous equipment.

    The following is required miscellaneous equipment:
    (a) [Reserved]
    (b) Two or more independent sources of electrical energy.
    (c) Electrical protective devices, as prescribed in this part.
    (d) Two systems for two-way radio communications, with controls for 
each accessible from each pilot station, designed and installed so that 
failure of one system will not preclude operation of the other system. 
The use of a common antenna system is acceptable if adequate reliability 
is shown.
    (e) Two systems for radio navigation, with controls for each 
accessible from each pilot station, designed and installed so that 
failure of one system will not preclude operation of the other system. 
The use of a common antenna system is acceptable if adequate reliability 
is shown.

[Amdt. 25-23, 35 FR 5678, Apr. 8, 1970, as amended by Amdt. 25-46, 43 FR 
50598, Oct. 30, 1978; Amdt. 25-54, 45 FR 60173, Sept. 11, 1980; Amdt. 
25-72, 55 FR 29785, July 20, 1990]



Sec. 25.1309  Equipment, systems, and installations.

    (a) The equipment, systems, and installations whose functioning is 
required by this subchapter, must be designed to ensure that they 
perform their intended functions under any foreseeable operating 
condition.
    (b) The airplane systems and associated components, considered 
separately and in relation to other systems, must be designed so that--
    (1) The occurrence of any failure condition which would prevent the 
continued safe flight and landing of the airplane is extremely 
improbable, and
    (2) The occurrence of any other failure conditions which would 
reduce the capability of the airplane or the ability of the crew to cope 
with adverse operating conditions is improbable.
    (c) Warning information must be provided to alert the crew to unsafe 
system operating conditions, and to enable them to take appropriate 
corrective action. Systems, controls, and associated monitoring and 
warning means must be designed to minimize crew errors which could 
create additional hazards.
    (d) Compliance with the requirements of paragraph (b) of this 
section must be shown by analysis, and where necessary, by appropriate 
ground, flight, or simulator tests. The analysis must consider--
    (1) Possible modes of failure, including malfunctions and damage 
from external sources.
    (2) The probability of multiple failures and undetected failures.
    (3) The resulting effects on the airplane and occupants, considering 
the stage of flight and operating conditions, and
    (4) The crew warning cues, corrective action required, and the 
capability of detecting faults.
    (e) In showing compliance with paragraphs (a) and (b) of this 
section with regard to the electrical system and equipment design and 
installation, critical environmental conditions must be considered. For 
electrical generation, distribution, and utilization equipment required 
by or used in complying with this chapter, except equipment covered by 
Technical Standard Orders containing environmental test procedures, the 
ability to provide continuous, safe service under foreseeable 
environmental conditions may be shown by environmental tests, design 
analysis, or reference to previous comparable service experience on 
other aircraft.

[[Page 509]]

    (f) EWIS must be assessed in accordance with the requirements of 
Sec. 25.1709.

[Amdt. 25-23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25-38, 41 FR 
55467, Dec. 20, 1976; Amdt. 25-41, 42 FR 36970, July 18, 1977; Amdt. 25-
123, 72 FR 63405, Nov. 8, 2007]



Sec. 25.1310  Power source capacity and distribution.

    (a) Each installation whose functioning is required for type 
certification or under operating rules and that requires a power supply 
is an ``essential load'' on the power supply. The power sources and the 
system must be able to supply the following power loads in probable 
operating combinations and for probable durations:
    (1) Loads connected to the system with the system functioning 
normally.
    (2) Essential loads, after failure of any one prime mover, power 
converter, or energy storage device.
    (3) Essential loads after failure of--
    (i) Any one engine on two-engine airplanes; and
    (ii) Any two engines on airplanes with three or more engines.
    (4) Essential loads for which an alternate source of power is 
required, after any failure or malfunction in any one power supply 
system, distribution system, or other utilization system.
    (b) In determining compliance with paragraphs (a)(2) and (3) of this 
section, the power loads may be assumed to be reduced under a monitoring 
procedure consistent with safety in the kinds of operation authorized. 
Loads not required in controlled flight need not be considered for the 
two-engine-inoperative condition on airplanes with three or more 
engines.

[Amdt. 25-123, 72 FR 63405, Nov. 8, 2007]



Sec. 25.1316  Electrical and electronic system lightning protection.

    (a) Each electrical and electronic system that performs a function, 
for which failure would prevent the continued safe flight and landing of 
the airplane, must be designed and installed so that--
    (1) The function is not adversely affected during and after the time 
the airplane is exposed to lightning; and
    (2) The system automatically recovers normal operation of that 
function in a timely manner after the airplane is exposed to lightning.
    (b) Each electrical and electronic system that performs a function, 
for which failure would reduce the capability of the airplane or the 
ability of the flightcrew to respond to an adverse operating condition, 
must be designed and installed so that the function recovers normal 
operation in a timely manner after the airplane is exposed to lightning.

[Doc. No. FAA-2010-0224, Amdt. 25-134, 76 FR 33135, June 8, 2011]



Sec. 25.1317  High-intensity Radiated Fields (HIRF) Protection.

    (a) Except as provided in paragraph (d) of this section, each 
electrical and electronic system that performs a function whose failure 
would prevent the continued safe flight and landing of the airplane must 
be designed and installed so that--
    (1) The function is not adversely affected during and after the time 
the airplane is exposed to HIRF environment I, as described in appendix 
L to this part;
    (2) The system automatically recovers normal operation of that 
function, in a timely manner, after the airplane is exposed to HIRF 
environment I, as described in appendix L to this part, unless the 
system's recovery conflicts with other operational or functional 
requirements of the system; and
    (3) The system is not adversely affected during and after the time 
the airplane is exposed to HIRF environment II, as described in appendix 
L to this part.
    (b) Each electrical and electronic system that performs a function 
whose failure would significantly reduce the capability of the airplane 
or the ability of the flightcrew to respond to an adverse operating 
condition must be designed and installed so the system is not adversely 
affected when the equipment providing these functions is exposed to 
equipment HIRF test level 1 or 2, as described in appendix L to this 
part.
    (c) Each electrical and electronic system that performs a function 
whose failure would reduce the capability of the airplane or the ability 
of the

[[Page 510]]

flightcrew to respond to an adverse operating condition must be designed 
and installed so the system is not adversely affected when the equipment 
providing the function is exposed to equipment HIRF test level 3, as 
described in appendix L to this part.
    (d) Before December 1, 2012, an electrical or electronic system that 
performs a function whose failure would prevent the continued safe 
flight and landing of an airplane may be designed and installed without 
meeting the provisions of paragraph (a) provided--
    (1) The system has previously been shown to comply with special 
conditions for HIRF, prescribed under Sec. 21.16, issued before 
December 1, 2007;
    (2) The HIRF immunity characteristics of the system have not changed 
since compliance with the special conditions was demonstrated; and
    (3) The data used to demonstrate compliance with the special 
conditions is provided.

[Doc. No. FAA-2006-23657, 72 FR 44025, Aug. 6, 2007]

                        Instruments: Installation



Sec. 25.1321  Arrangement and visibility.

    (a) Each flight, navigation, and powerplant instrument for use by 
any pilot must be plainly visible to him from his station with the 
minimum practicable deviation from his normal position and line of 
vision when he is looking forward along the flight path.
    (b) The flight instruments required by Sec. 25.1303 must be grouped 
on the instrument panel and centered as nearly as practicable about the 
vertical plane of the pilot's forward vision. In addition--
    (1) The instrument that most effectively indicates attitude must be 
on the panel in the top center position;
    (2) The instrument that most effectively indicates airspeed must be 
adjacent to and directly to the left of the instrument in the top center 
position:
    (3) The instrument that most effectively indicates altitude must be 
adjacent to and directly to the right of the instrument in the top 
center position; and
    (4) The instrument that most effectively indicates direction of 
flight must be adjacent to and directly below the instrument in the top 
center position.
    (c) Required powerplant instruments must be closely grouped on the 
instrument panel. In addition--
    (1) The location of identical powerplant instruments for the engines 
must prevent confusion as to which engine each instrument relates; and
    (2) Powerplant instruments vital to the safe operation of the 
airplane must be plainly visible to the appropriate crewmembers.
    (d) Instrument panel vibration may not damage or impair the accuracy 
of any instrument.
    (e) If a visual indicator is provided to indicate malfunction of an 
instrument, it must be effective under all probable cockpit lighting 
conditions.

[Amdt. 25-23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25-41, 42 FR 
36970, July 18, 1977]



Sec. 25.1322  Flightcrew alerting.

    (a) Flightcrew alerts must:
    (1) Provide the flightcrew with the information needed to:
    (i) Identify non-normal operation or airplane system conditions, and
    (ii) Determine the appropriate actions, if any.
    (2) Be readily and easily detectable and intelligible by the 
flightcrew under all foreseeable operating conditions, including 
conditions where multiple alerts are provided.
    (3) Be removed when the alerting condition no longer exists.
    (b) Alerts must conform to the following prioritization hierarchy 
based on the urgency of flightcrew awareness and response.
    (1) Warning: For conditions that require immediate flightcrew 
awareness and immediate flightcrew response.
    (2) Caution: For conditions that require immediate flightcrew 
awareness and subsequent flightcrew response.
    (3) Advisory: For conditions that require flightcrew awareness and 
may require subsequent flightcrew response.
    (c) Warning and caution alerts must:
    (1) Be prioritized within each category, when necessary.
    (2) Provide timely attention-getting cues through at least two 
different senses by a combination of aural, visual, or tactile 
indications.

[[Page 511]]

    (3) Permit each occurrence of the attention-getting cues required by 
paragraph (c)(2) of this section to be acknowledged and suppressed, 
unless they are required to be continuous.
    (d) The alert function must be designed to minimize the effects of 
false and nuisance alerts. In particular, it must be designed to:
    (1) Prevent the presentation of an alert that is inappropriate or 
unnecessary.
    (2) Provide a means to suppress an attention-getting component of an 
alert caused by a failure of the alerting function that interferes with 
the flightcrew's ability to safely operate the airplane. This means must 
not be readily available to the flightcrew so that it could be operated 
inadvertently or by habitual reflexive action. When an alert is 
suppressed, there must be a clear and unmistakable annunciation to the 
flightcrew that the alert has been suppressed.
    (e) Visual alert indications must:
    (1) Conform to the following color convention:
    (i) Red for warning alert indications.
    (ii) Amber or yellow for caution alert indications.
    (iii) Any color except red or green for advisory alert indications.
    (2) Use visual coding techniques, together with other alerting 
function elements on the flight deck, to distinguish between warning, 
caution, and advisory alert indications, if they are presented on 
monochromatic displays that are not capable of conforming to the color 
convention in paragraph (e)(1) of this section.
    (f) Use of the colors red, amber, and yellow on the flight deck for 
functions other than flightcrew alerting must be limited and must not 
adversely affect flightcrew alerting.

[Amdt. 25-131, 75 FR 67209, Nov. 2, 2010]



Sec. 25.1323  Airspeed indicating system.

    For each airspeed indicating system, the following apply:
    (a) Each airspeed indicating instrument must be approved and must be 
calibrated to indicate true airspeed (at sea level with a standard 
atmosphere) with a minimum practicable instrument calibration error when 
the corresponding pitot and static pressures are applied.
    (b) Each system must be calibrated to determine the system error 
(that is, the relation between IAS and CAS) in flight and during the 
accelerated takeoff ground run. The ground run calibration must be 
determined--
    (1) From 0.8 of the minimum value of V1 to the maximum 
value of V2, considering the approved ranges of altitude and 
weight; and
    (2) With the flaps and power settings corresponding to the values 
determined in the establishment of the takeoff path under Sec. 25.111 
assuming that the critical engine fails at the minimum value of 
V1.
    (c) The airspeed error of the installation, excluding the airspeed 
indicator instrument calibration error, may not exceed three percent or 
five knots, whichever is greater, throughout the speed range, from--
    (1) VMO to 1.23 VSR1, with flaps retracted; and
    (2) 1.23 VSR0 to VFE with flaps in the landing 
position.
    (d) From 1.23 VSR to the speed at which stall warning 
begins, the IAS must change perceptibly with CAS and in the same sense, 
and at speeds below stall warning speed the IAS must not change in an 
incorrect sense.
    (e) From VMO to VMO + \2/3\ (VDF - 
VMO), the IAS must change perceptibly with CAS and in the 
same sense, and at higher speeds up to VDF the IAS must not 
change in an incorrect sense.
    (f) There must be no indication of airspeed that would cause undue 
difficulty to the pilot during the takeoff between the initiation of 
rotation and the achievement of a steady climbing condition.
    (g) The effects of airspeed indicating system lag may not introduce 
significant takeoff indicated airspeed bias, or significant errors in 
takeoff or accelerate-stop distances.
    (h) Each system must be arranged, so far as practicable, to prevent 
malfunction or serious error due to the entry of moisture, dirt, or 
other substances.
    (i) Each system must have a heated pitot tube or an equivalent means 
of preventing malfunction due to icing.
    (j) Where duplicate airspeed indicators are required, their 
respective pitot

[[Page 512]]

tubes must be far enough apart to avoid damage to both tubes in a 
collision with a bird.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 
49 FR 6849, Feb. 23, 1984; Amdt. 25-108, 67 FR 70828, Nov. 26, 2002; 
Amdt. 25-109, 67 FR 76656, Dec. 12, 2002]

    Effective Date Note: At 79 FR 65527, Nov. 4, 2014, Sec. 25.1323 was 
amended by revising paragraph (i), effective Jan. 5, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.1323  Airspeed indicating system.

                                * * * * *

    (i) Each system must have a heated pitot tube or an equivalent means 
of preventing malfunction in the heavy rain conditions defined in Table 
1 of this section; mixed phase and ice crystal conditions as defined in 
part 33, Appendix D, of this chapter; the icing conditions defined in 
Appendix C of this part; and the following icing conditions specified in 
Appendix O of this part:
    (1) For airplanes certificated in accordance with Sec. 
25.1420(a)(1), the icing conditions that the airplane is certified to 
safely exit following detection.
    (2) For airplanes certificated in accordance with Sec. 
25.1420(a)(2), the icing conditions that the airplane is certified to 
safely operate in and the icing conditions that the airplane is 
certified to safely exit following detection.
    (3) For airplanes certificated in accordance with Sec. 
25.1420(a)(3) and for airplanes not subject to Sec. 25.1420, all icing 
conditions.

                       Table 1--Heavy Rain Conditions for Airspeed Indicating System Tests
----------------------------------------------------------------------------------------------------------------
                 Altitude range                   Liquid water          Horizontal extent           Droplet MVD
------------------------------------------------     content    ------------------------------------------------
                                                ----------------
             (ft)                     (m)            (g/m3)           (km)           (nmiles)       ([micro]m)
----------------------------------------------------------------------------------------------------------------
0 to 10 000..................  0 to 3000.......               1             100             50              1000
                                                              6               5              3              2000
                                                             15               1              0.5            2000
----------------------------------------------------------------------------------------------------------------

                                * * * * *



Sec. 25.1324  Angle of attack system.

    Each angle of attack system sensor must be heated or have an 
equivalent means of preventing malfunction in the heavy rain conditions 
defined in Table 1 of Sec. 25.1323, the mixed phase and ice crystal 
conditions as defined in part 33, Appendix D, of this chapter, the icing 
conditions defined in Appendix C of this part, and the following icing 
conditions specified in Appendix O of this part:
    (a) For airplanes certificated in accordance with Sec. 
25.1420(a)(1), the icing conditions that the airplane is certified to 
safely exit following detection.
    (b) For airplanes certificated in accordance with Sec. 
25.1420(a)(2), the icing conditions that the airplane is certified to 
safely operate in and the icing conditions that the airplane is 
certified to safely exit following detection.
    (c) For airplanes certificated in accordance with Sec. 
25.1420(a)(3) and for airplanes not subject to Sec. 25.1420, all icing 
conditions.

[79 FR 65527, Nov. 4, 2014]

    Effective Date Note: At 79 FR 65527, Nov. 4, 2014, Sec. 25.1324 was 
added, effective Jan. 5, 2015.



Sec. 25.1325  Static pressure systems.

    (a) Each instrument with static air case connections must be vented 
to the outside atmosphere through an appropriate piping system.
    (b) Each static port must be designed and located in such manner 
that the static pressure system performance is least affected by airflow 
variation, or by moisture or other foreign matter, and that the 
correlation between air pressure in the static pressure system and true 
ambient atmospheric static pressure is not changed when the airplane is 
exposed to the continuous and intermittent maximum icing conditions 
defined in appendix C of this part.
    (c) The design and installation of the static pressure system must 
be such that--
    (1) Positive drainage of moisture is provided; chafing of the tubing 
and excessive distortion or restriction at bends in the tubing is 
avoided; and the materials used are durable, suitable for the purpose 
intended, and protected against corrosion; and

[[Page 513]]

    (2) It is airtight except for the port into the atmosphere. A proof 
test must be conducted to demonstrate the integrity of the static 
pressure system in the following manner:
    (i) Unpressurized airplanes. Evacuate the static pressure system to 
a pressure differential of approximately 1 inch of mercury or to a 
reading on the altimeter, 1,000 feet above the airplane elevation at the 
time of the test. Without additional pumping for a period of 1 minute, 
the loss of indicated altitude must not exceed 100 feet on the 
altimeter.
    (ii) Pressurized airplanes. Evacuate the static pressure system 
until a pressure differential equivalent to the maximum cabin pressure 
differential for which the airplane is type certificated is achieved. 
Without additional pumping for a period of 1 minute, the loss of 
indicated altitude must not exceed 2 percent of the equivalent altitude 
of the maximum cabin differential pressure or 100 feet, whichever is 
greater.
    (d) Each pressure altimeter must be approved and must be calibrated 
to indicate pressure altitude in a standard atmosphere, with a minimum 
practicable calibration error when the corresponding static pressures 
are applied.
    (e) Each system must be designed and installed so that the error in 
indicated pressure altitude, at sea level, with a standard atmosphere, 
excluding instrument calibration error, does not result in an error of 
more than 30 feet per 100 knots speed for the 
appropriate configuration in the speed range between 1.23 
VSR0 with flaps extended and 1.7 VSR1 with flaps 
retracted. However, the error need not be less than 30 feet.
    (f) If an altimeter system is fitted with a device that provides 
corrections to the altimeter indication, the device must be designed and 
installed in such manner that it can be bypassed when it malfunctions, 
unless an alternate altimeter system is provided. Each correction device 
must be fitted with a means for indicating the occurrence of reasonably 
probable malfunctions, including power failure, to the flight crew. The 
indicating means must be effective for any cockpit lighting condition 
likely to occur.
    (g) Except as provided in paragraph (h) of this section, if the 
static pressure system incorporates both a primary and an alternate 
static pressure source, the means for selecting one or the other source 
must be designed so that--
    (1) When either source is selected, the other is blocked off; and
    (2) Both sources cannot be blocked off simultaneously.
    (h) For unpressurized airplanes, paragraph (g)(1) of this section 
does not apply if it can be demonstrated that the static pressure system 
calibration, when either static pressure source is selected, is not 
changed by the other static pressure source being open or blocked.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-5, 30 
FR 8261, June 29, 1965; Amdt. 25-12, 32 FR 7587, May 24, 1967; Amdt. 25-
41, 42 FR 36970, July 18, 1977; Amdt. 25-108, 67 FR 70828, Nov. 26, 
2002]

    Effective Date Note: At 79 FR 65527, Nov. 4, 2014, Sec. 25.1325 was 
amended by revising paragraph (b), effective Jan. 5, 2015. For the 
convenience of the user, the revised text is set forth as follows:



Sec. 25.1325  Static pressure systems.

                                * * * * *

    (b) Each static port must be designed and located so that:
    (1) The static pressure system performance is least affected by 
airflow variation, or by moisture or other foreign matter; and
    (2) The correlation between air pressure in the static pressure 
system and true ambient atmospheric static pressure is not changed when 
the airplane is exposed to the icing conditions defined in Appendix C of 
this part, and the following icing conditions specified in Appendix O of 
this part:
    (i) For airplanes certificated in accordance with Sec. 
25.1420(a)(1), the icing conditions that the airplane is certified to 
safely exit following detection.
    (ii) For airplanes certificated in accordance with Sec. 
25.1420(a)(2), the icing conditions that the airplane is certified to 
safely operate in and the icing conditions that the airplane is 
certified to safely exit following detection.
    (iii) For airplanes certificated in accordance with Sec. 
25.1420(a)(3) and for airplanes not subject to Sec. 25.1420, all icing 
conditions.

                                * * * * *

[[Page 514]]



Sec. 25.1326  Pitot heat indication systems.

    If a flight instrument pitot heating system is installed, an 
indication system must be provided to indicate to the flight crew when 
that pitot heating system is not operating. The indication system must 
comply with the following requirements:
    (a) The indication provided must incorporate an amber light that is 
in clear view of a flight crewmember.
    (b) The indication provided must be designed to alert the flight 
crew if either of the following conditions exist:
    (1) The pitot heating system is switched ``off''.
    (2) The pitot heating system is switched ``on'' and any pitot tube 
heating element is inoperative.

[Amdt. 25-43, 43 FR 10339, Mar. 13, 1978]



Sec. 25.1327  Magnetic direction indicator.

    (a) Each magnetic direction indicator must be installed so that its 
accuracy is not excessively affected by the airplane's vibration or 
magnetic fields.
    (b) The compensated installation may not have a deviation, in level 
flight, greater than 10 degrees on any heading.



Sec. 25.1329  Flight guidance system.

    (a) Quick disengagement controls for the autopilot and autothrust 
functions must be provided for each pilot. The autopilot quick 
disengagement controls must be located on both control wheels (or 
equivalent). The autothrust quick disengagement controls must be located 
on the thrust control levers. Quick disengagement controls must be 
readily accessible to each pilot while operating the control wheel (or 
equivalent) and thrust control levers.
    (b) The effects of a failure of the system to disengage the 
autopilot or autothrust functions when manually commanded by the pilot 
must be assessed in accordance with the requirements of Sec. 25.1309.
    (c) Engagement or switching of the flight guidance system, a mode, 
or a sensor may not cause a transient response of the airplane's flight 
path any greater than a minor transient, as defined in paragraph (n)(1) 
of this section.
    (d) Under normal conditions, the disengagement of any automatic 
control function of a flight guidance system may not cause a transient 
response of the airplane's flight path any greater than a minor 
transient.
    (e) Under rare normal and non-normal conditions, disengagement of 
any automatic control function of a flight guidance system may not 
result in a transient any greater than a significant transient, as 
defined in paragraph (n)(2) of this section.
    (f) The function and direction of motion of each command reference 
control, such as heading select or vertical speed, must be plainly 
indicated on, or adjacent to, each control if necessary to prevent 
inappropriate use or confusion.
    (g) Under any condition of flight appropriate to its use, the flight 
guidance system may not produce hazardous loads on the airplane, nor 
create hazardous deviations in the flight path. This applies to both 
fault-free operation and in the event of a malfunction, and assumes that 
the pilot begins corrective action within a reasonable period of time.
    (h) When the flight guidance system is in use, a means must be 
provided to avoid excursions beyond an acceptable margin from the speed 
range of the normal flight envelope. If the airplane experiences an 
excursion outside this range, a means must be provided to prevent the 
flight guidance system from providing guidance or control to an unsafe 
speed.
    (i) The flight guidance system functions, controls, indications, and 
alerts must be designed to minimize flightcrew errors and confusion 
concerning the behavior and operation of the flight guidance system. 
Means must be provided to indicate the current mode of operation, 
including any armed modes, transitions, and reversions. Selector switch 
position is not an acceptable means of indication. The controls and 
indications must be grouped and presented in a logical and consistent 
manner. The indications must be visible to each pilot under all expected 
lighting conditions.
    (j) Following disengagement of the autopilot, a warning (visual and 
auditory) must be provided to each pilot and be timely and distinct from 
all other cockpit warnings.

[[Page 515]]

    (k) Following disengagement of the autothrust function, a caution 
must be provided to each pilot.
    (l) The autopilot may not create a potential hazard when the 
flightcrew applies an override force to the flight controls.
    (m) During autothrust operation, it must be possible for the 
flightcrew to move the thrust levers without requiring excessive force. 
The autothrust may not create a potential hazard when the flightcrew 
applies an override force to the thrust levers.
    (n) For purposes of this section, a transient is a disturbance in 
the control or flight path of the airplane that is not consistent with 
response to flightcrew inputs or environmental conditions.
    (1) A minor transient would not significantly reduce safety margins 
and would involve flightcrew actions that are well within their 
capabilities. A minor transient may involve a slight increase in 
flightcrew workload or some physical discomfort to passengers or cabin 
crew.
    (2) A significant transient may lead to a significant reduction in 
safety margins, an increase in flightcrew workload, discomfort to the 
flightcrew, or physical distress to the passengers or cabin crew, 
possibly including non-fatal injuries. Significant transients do not 
require, in order to remain within or recover to the normal flight 
envelope, any of the following:
    (i) Exceptional piloting skill, alertness, or strength.
    (ii) Forces applied by the pilot which are greater than those 
specified in Sec. 25.143(c).
    (iii) Accelerations or attitudes in the airplane that might result 
in further hazard to secured or non-secured occupants.

[Doc. No. FAA-2004-18775, 71 FR 18191, Apr. 11, 2006]



Sec. 25.1331  Instruments using a power supply.

    (a) For each instrument required by Sec. 25.1303(b) that uses a 
power supply, the following apply:
    (1) Each instrument must have a visual means integral with, the 
instrument, to indicate when power adequate to sustain proper instrument 
performance is not being supplied. The power must be measured at or near 
the point where it enters the instruments. For electric instruments, the 
power is considered to be adequate when the voltage is within approved 
limits.
    (2) Each instrument must, in the event of the failure of one power 
source, be supplied by another power source. This may be accomplished 
automatically or by manual means.
    (3) If an instrument presenting navigation data receives information 
from sources external to that instrument and loss of that information 
would render the presented data unreliable, the instrument must 
incorporate a visual means to warn the crew, when such loss of 
information occurs, that the presented data should not be relied upon.
    (b) As used in this section, ``instrument'' includes devices that 
are physically contained in one unit, and devices that are composed of 
two or more physically separate units or components connected together 
(such as a remote indicating gyroscopic direction indicator that 
includes a magnetic sensing element, a gyroscopic unit, an amplifier and 
an indicator connected together).

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 
42 FR 36970, July 18, 1977]



Sec. 25.1333  Instrument systems.

    For systems that operate the instruments required by Sec. 
25.1303(b) which are located at each pilot's station--
    (a) Means must be provided to connect the required instruments at 
the first pilot's station to operating systems which are independent of 
the operating systems at other flight crew stations, or other equipment;
    (b) The equipment, systems, and installations must be designed so 
that one display of the information essential to the safety of flight 
which is provided by the instruments, including attitude, direction, 
airspeed, and altitude will remain available to the pilots, without 
additional crewmember action, after any single failure or combination of 
failures that is not shown to be extremely improbable; and
    (c) Additional instruments, systems, or equipment may not be 
connected to

[[Page 516]]

the operating systems for the required instruments, unless provisions 
are made to ensure the continued normal functioning of the required 
instruments in the event of any malfunction of the additional 
instruments, systems, or equipment which is not shown to be extremely 
improbable.

[Amdt. 25-23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25-41, 42 FR 
36970, July 18, 1977]



Sec. 25.1337  Powerplant instruments.

    (a) Instruments and instrument lines. (1) Each powerplant and 
auxiliary power unit instrument line must meet the requirements of 
Sec. Sec. 25.993 and 25.1183.
    (2) Each line carrying flammable fluids under pressure must--
    (i) Have restricting orifices or other safety devices at the source 
of pressure to prevent the escape of excessive fluid if the line fails; 
and
    (ii) Be installed and located so that the escape of fluids would not 
create a hazard.
    (3) Each powerplant and auxiliary power unit instrument that 
utilizes flammable fluids must be installed and located so that the 
escape of fluid would not create a hazard.
    (b) Fuel quantity indicator. There must be means to indicate to the 
flight crewmembers, the quantity, in gallons or equivalent units, of 
usable fuel in each tank during flight. In addition--
    (1) Each fuel quantity indicator must be calibrated to read ``zero'' 
during level flight when the quantity of fuel remaining in the tank is 
equal to the unusable fuel supply determined under Sec. 25.959;
    (2) Tanks with interconnected outlets and airspaces may be treated 
as one tank and need not have separate indicators; and
    (3) Each exposed sight gauge, used as a fuel quantity indicator, 
must be protected against damage.
    (c) Fuel flowmeter system. If a fuel flowmeter system is installed, 
each metering component must have a means for bypassing the fuel supply 
if malfunction of that component severely restricts fuel flow.
    (d) Oil quantity indicator. There must be a stick gauge or 
equivalent means to indicate the quantity of oil in each tank. If an oil 
transfer or reserve oil supply system is installed, there must be a 
means to indicate to the flight crew, in flight, the quantity of oil in 
each tank.
    (e) Turbopropeller blade position indicator. Required turbopropeller 
blade position indicators must begin indicating before the blade moves 
more than eight degrees below the flight low pitch stop. The source of 
indication must directly sense the blade position.
    (f) Fuel pressure indicator. There must be means to measure fuel 
pressure, in each system supplying reciprocating engines, at a point 
downstream of any fuel pump except fuel injection pumps. In addition--
    (1) If necessary for the maintenance of proper fuel delivery 
pressure, there must be a connection to transmit the carburetor air 
intake static pressure to the proper pump relief valve connection; and
    (2) If a connection is required under paragraph (f)(1) of this 
section, the gauge balance lines must be independently connected to the 
carburetor inlet pressure to avoid erroneous readings.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 
42 FR 15044, Mar. 17, 1977]

                    Electrical Systems and Equipment



Sec. 25.1351  General.

    (a) Electrical system capacity. The required generating capacity, 
and number and kinds of power sources must--
    (1) Be determined by an electrical load analysis; and
    (2) Meet the requirements of Sec. 25.1309.
    (b) Generating system. The generating system includes electrical 
power sources, main power busses, transmission cables, and associated 
control, regulation, and protective devices. It must be designed so 
that--
    (1) Power sources function properly when independent and when 
connected in combination;
    (2) No failure or malfunction of any power source can create a 
hazard or impair the ability of remaining sources to supply essential 
loads;
    (3) The system voltage and frequency (as applicable) at the 
terminals of all essential load equipment can be maintained within the 
limits for which the

[[Page 517]]

equipment is designed, during any probable operating condition; and
    (4) System transients due to switching, fault clearing, or other 
causes do not make essential loads inoperative, and do not cause a smoke 
or fire hazard.
    (5) There are means accessible, in flight, to appropriate 
crewmembers for the individual and collective disconnection of the 
electrical power sources from the system.
    (6) There are means to indicate to appropriate crewmembers the 
generating system quantities essential for the safe operation of the 
system, such as the voltage and current supplied by each generator.
    (c) External power. If provisions are made for connecting external 
power to the airplane, and that external power can be electrically 
connected to equipment other than that used for engine starting, means 
must be provided to ensure that no external power supply having a 
reverse polarity, or a reverse phase sequence, can supply power to the 
airplane's electrical system.
    (d) Operation without normal electrical power. It must be shown by 
analysis, tests, or both, that the airplane can be operated safely in 
VFR conditions, for a period of not less than five minutes, with the 
normal electrical power (electrical power sources excluding the battery) 
inoperative, with critical type fuel (from the standpoint of flameout 
and restart capability), and with the airplane initially at the maximum 
certificated altitude. Parts of the electrical system may remain on if--
    (1) A single malfunction, including a wire bundle or junction box 
fire, cannot result in loss of both the part turned off and the part 
turned on; and
    (2) The parts turned on are electrically and mechanically isolated 
from the parts turned off.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 
42 FR 36970, July 18, 1977; Amdt. 25-72, 55 FR 29785, July 20, 1990]



Sec. 25.1353  Electrical equipment and installations.

    (a) Electrical equipment and controls must be installed so that 
operation of any one unit or system of units will not adversely affect 
the simultaneous operation of any other electrical unit or system 
essential to safe operation. Any electrical interference likely to be 
present in the airplane must not result in hazardous effects on the 
airplane or its systems.
    (b) Storage batteries must be designed and installed as follows:
    (1) Safe cell temperatures and pressures must be maintained during 
any probable charging or discharging condition. No uncontrolled increase 
in cell temperature may result when the battery is recharged (after 
previous complete discharge)--
    (i) At maximum regulated voltage or power;
    (ii) During a flight of maximum duration; and
    (iii) Under the most adverse cooling condition likely to occur in 
service.
    (2) Compliance with paragraph (b)(1) of this section must be shown 
by test unless experience with similar batteries and installations has 
shown that maintaining safe cell temperatures and pressures presents no 
problem.
    (3) No explosive or toxic gases emitted by any battery in normal 
operation, or as the result of any probable malfunction in the charging 
system or battery installation, may accumulate in hazardous quantities 
within the airplane.
    (4) No corrosive fluids or gases that may escape from the battery 
may damage surrounding airplane structures or adjacent essential 
equipment.
    (5) Each nickel cadmium battery installation must have provisions to 
prevent any hazardous effect on structure or essential systems that may 
be caused by the maximum amount of heat the battery can generate during 
a short circuit of the battery or of individual cells.
    (6) Nickel cadmium battery installations must have--
    (i) A system to control the charging rate of the battery 
automatically so as to prevent battery overheating;
    (ii) A battery temperature sensing and over-temperature warning 
system with a means for disconnecting the battery from its charging 
source in the event of an over-temperature condition; or

[[Page 518]]

    (iii) A battery failure sensing and warning system with a means for 
disconnecting the battery from its charging source in the event of 
battery failure.
    (c) Electrical bonding must provide an adequate electrical return 
path under both normal and fault conditions, on airplanes having 
grounded electrical systems.

[Amdt. 25-123, 72 FR 63405, Nov. 8, 2007]



Sec. 25.1355  Distribution system.

    (a) The distribution system includes the distribution busses, their 
associated feeders, and each control and protective device.
    (b) [Reserved]
    (c) If two independent sources of electrical power for particular 
equipment or systems are required by this chapter, in the event of the 
failure of one power source for such equipment or system, another power 
source (including its separate feeder) must be automatically provided or 
be manually selectable to maintain equipment or system operation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 
35 FR 5679, Apr. 8, 1970; Amdt. 25-38, 41 FR 55468, Dec. 20, 1976]



Sec. 25.1357  Circuit protective devices.

    (a) Automatic protective devices must be used to minimize distress 
to the electrical system and hazard to the airplane in the event of 
wiring faults or serious malfunction of the system or connected 
equipment.
    (b) The protective and control devices in the generating system must 
be designed to de-energize and disconnect faulty power sources and power 
transmission equipment from their associated busses with sufficient 
rapidity to provide protection from hazardous over-voltage and other 
malfunctioning.
    (c) Each resettable circuit protective device must be designed so 
that, when an overload or circuit fault exists, it will open the circuit 
irrespective of the position of the operating control.
    (d) If the ability to reset a circuit breaker or replace a fuse is 
essential to safety in flight, that circuit breaker or fuse must be 
located and identified so that it can be readily reset or replaced in 
flight. Where fuses are used, there must be spare fuses for use in 
flight equal to at least 50% of the number of fuses of each rating 
required for complete circuit protection.
    (e) Each circuit for essential loads must have individual circuit 
protection. However, individual protection for each circuit in an 
essential load system (such as each position light circuit in a system) 
is not required.
    (f) For airplane systems for which the ability to remove or reset 
power during normal operations is necessary, the system must be designed 
so that circuit breakers are not the primary means to remove or reset 
system power unless specifically designed for use as a switch.
    (g) Automatic reset circuit breakers may be used as integral 
protectors for electrical equipment (such as thermal cut-outs) if there 
is circuit protection to protect the cable to the equipment.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-123, 
72 FR 63405, Nov. 8, 2007]



Sec. 25.1360  Precautions against injury.

    (a) Shock. The electrical system must be designed to minimize risk 
of electric shock to crew, passengers, and servicing personnel and to 
maintenance personnel using normal precautions.
    (b) Burns. The temperature of any part that may be handled by a 
crewmember during normal operations must not cause dangerous inadvertent 
movement by the crewmember or injury to the crewmember.

[Amdt. 25-123, 72 FR 63406, Nov. 8, 2007]



Sec. 25.1362  Electrical supplies for emergency conditions.

    A suitable electrical supply must be provided to those services 
required for emergency procedures after an emergency landing or 
ditching. The circuits for these services must be designed, protected, 
and installed so that the risk of the services being rendered 
ineffective under these emergency conditions is minimized.

[Amdt. 25-123, 72 FR 63406, Nov. 8, 2007]



Sec. 25.1363  Electrical system tests.

    (a) When laboratory tests of the electrical system are conducted--

[[Page 519]]

    (1) The tests must be performed on a mock-up using the same 
generating equipment used in the airplane;
    (2) The equipment must simulate the electrical characteristics of 
the distribution wiring and connected loads to the extent necessary for 
valid test results; and
    (3) Laboratory generator drives must simulate the actual prime 
movers on the airplane with respect to their reaction to generator 
loading, including loading due to faults.
    (b) For each flight condition that cannot be simulated adequately in 
the laboratory or by ground tests on the airplane, flight tests must be 
made.



Sec. 25.1365  Electrical appliances, motors, and transformers.

    (a) Domestic appliances must be designed and installed so that in 
the event of failures of the electrical supply or control system, the 
requirements of Sec. 25.1309(b), (c), and (d) will be satisfied. 
Domestic appliances are items such as cooktops, ovens, coffee makers, 
water heaters, refrigerators, and toilet flush systems that are placed 
on the airplane to provide service amenities to passengers.
    (b) Galleys and cooking appliances must be installed in a way that 
minimizes risk of overheat or fire.
    (c) Domestic appliances, particularly those in galley areas, must be 
installed or protected so as to prevent damage or contamination of other 
equipment or systems from fluids or vapors which may be present during 
normal operation or as a result of spillage, if such damage or 
contamination could create a hazardous condition.
    (d) Unless compliance with Sec. 25.1309(b) is provided by the 
circuit protective device required by Sec. 25.1357(a), electric motors 
and transformers, including those installed in domestic systems, must 
have a suitable thermal protection device to prevent overheating under 
normal operation and failure conditions, if overheating could create a 
smoke or fire hazard.

[Amdt. 25-123, 72 FR 63406, Nov. 8, 2007]

                                 Lights



Sec. 25.1381  Instrument lights.

    (a) The instrument lights must--
    (1) Provide sufficient illumination to make each instrument, switch 
and other device necessary for safe operation easily readable unless 
sufficient illumination is available from another source; and
    (2) Be installed so that--
    (i) Their direct rays are shielded from the pilot's eyes; and
    (ii) No objectionable reflections are visible to the pilot.
    (b) Unless undimmed instrument lights are satisfactory under each 
expected flight condition, there must be a means to control the 
intensity of illumination.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29785, July 20, 1990]



Sec. 25.1383  Landing lights.

    (a) Each landing light must be approved, and must be installed so 
that--
    (1) No objectionable glare is visible to the pilot;
    (2) The pilot is not adversely affected by halation; and
    (3) It provides enough light for night landing.
    (b) Except when one switch is used for the lights of a multiple 
light installation at one location, there must be a separate switch for 
each light.
    (c) There must be a means to indicate to the pilots when the landing 
lights are extended.



Sec. 25.1385  Position light system installation.

    (a) General. Each part of each position light system must meet the 
applicable requirements of this section and each system as a whole must 
meet the requirements of Sec. Sec. 25.1387 through 25.1397.
    (b) Forward position lights. Forward position lights must consist of 
a red and a green light spaced laterally as far apart as practicable and 
installed forward on the airplane so that, with the airplane in the 
normal flying position, the red light is on the left side and the green 
light is on the right side. Each light must be approved.
    (c) Rear position light. The rear position light must be a white 
light mounted as far aft as practicable on the tail or on each wing tip, 
and must be approved.

[[Page 520]]

    (d) Light covers and color filters. Each light cover or color filter 
must be at least flame resistant and may not change color or shape or 
lose any appreciable light transmission during normal use.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55468, Dec. 20, 1976]



Sec. 25.1387  Position light system dihedral angles.

    (a) Except as provided in paragraph (e) of this section, each 
forward and rear position light must, as installed, show unbroken light 
within the dihedral angles described in this section.
    (b) Dihedral angle L (left) is formed by two intersecting vertical 
planes, the first parallel to the longitudinal axis of the airplane, and 
the other at 110 degrees to the left of the first, as viewed when 
looking forward along the longitudinal axis.
    (c) Dihedral angle R (right) is formed by two intersecting vertical 
planes, the first parallel to the longitudinal axis of the airplane, and 
the other at 110 degrees to the right of the first, as viewed when 
looking forward along the longitudinal axis.
    (d) Dihedral angle A (aft) is formed by two intersecting vertical 
planes making angles of 70 degrees to the right and to the left, 
respectively, to a vertical plane passing through the longitudinal axis, 
as viewed when looking aft along the longitudinal axis.
    (e) If the rear position light, when mounted as far aft as 
practicable in accordance with Sec. 25.1385(c), cannot show unbroken 
light within dihedral angle A (as defined in paragraph (d) of this 
section), a solid angle or angles of obstructed visibility totaling not 
more than 0.04 steradians is allowable within that dihedral angle, if 
such solid angle is within a cone whose apex is at the rear position 
light and whose elements make an angle of 30[deg] with a vertical line 
passing through the rear position light.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-30, 
36 FR 21278, Nov. 5, 1971]



Sec. 25.1389  Position light distribution and intensities.

    (a) General. The intensities prescribed in this section must be 
provided by new equipment with light covers and color filters in place. 
Intensities must be determined with the light source operating at a 
steady value equal to the average luminous output of the source at the 
normal operating voltage of the airplane. The light distribution and 
intensity of each position light must meet the requirements of paragraph 
(b) of this section.
    (b) Forward and rear position lights. The light distribution and 
intensities of forward and rear position lights must be expressed in 
terms of minimum intensities in the horizontal plane, minimum 
intensities in any vertical plane, and maximum intensities in 
overlapping beams, within dihedral angles L, R, and A, and must meet the 
following requirements:
    (1) Intensities in the horizontal plane. Each intensity in the 
horizontal plane (the plane containing the longitudinal axis of the 
airplane and perpendicular to the plane of symmetry of the airplane) 
must equal or exceed the values in Sec. 25.1391.
    (2) Intensities in any vertical plane. Each intensity in any 
vertical plane (the plane perpendicular to the horizontal plane) must 
equal or exceed the appropriate value in Sec. 25.1393, where I is the 
minimum intensity prescribed in Sec. 25.1391 for the corresponding 
angles in the horizontal plane.
    (3) Intensities in overlaps between adjacent signals. No intensity 
in any overlap between adjacent signals may exceed the values given in 
Sec. 25.1395, except that higher intensities in overlaps may be used 
with main beam intensities substantially greater than the minima 
specified in Sec. Sec. 25.1391 and 25.1393 if the overlap intensities 
in relation to the main beam intensities do not adversely affect signal 
clarity. When the peak intensity of the forward position lights is more 
than 100 candles, the maximum overlap intensities between them may 
exceed the values given in Sec. 25.1395 if the overlap intensity in 
Area A is not more than 10 percent of peak position light intensity and 
the overlap intensity in Area B is not greater than 2.5

[[Page 521]]

percent of peak position light intensity.



Sec. 25.1391  Minimum intensities in the horizontal plane of forward
and rear position lights.

    Each position light intensity must equal or exceed the applicable 
values in the following table:

------------------------------------------------------------------------
                                        Angle from right or
                                       left of longitudinal    Intensity
   Dihedral angle (light included)      axis, measured from    (candles)
                                            dead ahead
------------------------------------------------------------------------
L and R (forward red and green).....  0[deg] to 10[deg].....          40
                                      10[deg] to 20[deg]....          30
                                      20[deg] to 110[deg]...           5
A (rear white)......................  110[deg] to 180[deg]..          20
------------------------------------------------------------------------



Sec. 25.1393  Minimum intensities in any vertical plane of forward
and rear position lights.

    Each position light intensity must equal or exceed the applicable 
values in the following table:

------------------------------------------------------------------------
                                                              Intensity,
          Angle above or below the horizontal plane                l
------------------------------------------------------------------------
0[deg]......................................................        1.00
0[deg] to 5[deg]............................................        0.90
5[deg] to 10[deg]...........................................        0.80
10[deg] to 15[deg]..........................................        0.70
15[deg] to 20[deg]..........................................        0.50
20[deg] to 30[deg]..........................................        0.30
30[deg] to 40[deg]..........................................        0.10
40[deg] to 90[deg]..........................................        0.05
------------------------------------------------------------------------



Sec. 25.1395  Maximum intensities in overlapping beams of forward 
and rear position lights.

    No position light intensity may exceed the applicable values in the 
following table, except as provided in Sec. 25.1389(b)(3).

------------------------------------------------------------------------
                                                     Maximum intensity
                                                 -----------------------
                    Overlaps                        Area A      Area B
                                                   (candles)   (candles)
------------------------------------------------------------------------
Green in dihedral angle L.......................          10           1
Red in dihedral angle R.........................          10           1
Green in dihedral angle A.......................           5           1
Red in dihedral angle A.........................           5           1
Rear white in dihedral angle L..................           5           1
Rear white in dihedral angle R..................           5           1
------------------------------------------------------------------------


Where--
    (a) Area A includes all directions in the adjacent dihedral angle 
that pass through the light source and intersect the common boundary 
plane at more than 10 degrees but less than 20 degrees; and
    (b) Area B includes all directions in the adjacent dihedral angle 
that pass through the light source and intersect the common boundary 
plane at more than 20 degrees.



Sec. 25.1397  Color specifications.

    Each position light color must have the applicable International 
Commission on Illumination chromaticity coordinates as follows:
    (a) Aviation red--

y is not greater than 0.335; and
z is not greater than 0.002.

    (b) Aviation green--

x is not greater than 0.440-0.320y ;
x is not greater than y-0.170; and
y is not less than 0.390-0.170x.

    (c) Aviation white--

x is not less than 0.300 and not greater than 0.540;
y is not less than x-0.040; or y0-0.010, whichever is the 
          smaller; and
y is not greater than x+0.020 nor 0.636-0.400x;
Where y0 is the y coordinate of the Planckian radiator for 
the value of x considered.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-27, 
36 FR 12972, July 10, 1971]



Sec. 25.1399  Riding light.

    (a) Each riding (anchor) light required for a seaplane or amphibian 
must be installed so that it can--
    (1) Show a white light for at least 2 nautical miles at night under 
clear atmospheric conditions; and
    (2) Show the maximum unbroken light practicable when the airplane is 
moored or drifting on the water.
    (b) Externally hung lights may be used.



Sec. 25.1401  Anticollision light system.

    (a) General. The airplane must have an anticollision light system 
that--
    (1) Consists of one or more approved anticollision lights located so 
that their light will not impair the crew's vision or detract from the 
conspicuity of the position lights; and
    (2) Meets the requirements of paragraphs (b) through (f) of this 
section.
    (b) Field of coverage. The system must consist of enough lights to 
illuminate the vital areas around the airplane considering the physical 
configuration

[[Page 522]]

and flight characteristics of the airplane. The field of coverage must 
extend in each direction within at least 75 degrees above and 75 degrees 
below the horizontal plane of the airplane, except that a solid angle or 
angles of obstructed visibility totaling not more than 0.03 steradians 
is allowable within a solid angle equal to 0.15 steradians centered 
about the longitudinal axis in the rearward direction.
    (c) Flashing characteristics. The arrangement of the system, that 
is, the number of light sources, beam width, speed of rotation, and 
other characteristics, must give an effective flash frequency of not 
less than 40, nor more than 100 cycles per minute. The effective flash 
frequency is the frequency at which the airplane's complete 
anticollision light system is observed from a distance, and applies to 
each sector of light including any overlaps that exist when the system 
consists of more than one light source. In overlaps, flash frequencies 
may exceed 100, but not 180 cycles per minute.
    (d) Color. Each anticollision light must be either aviation red or 
aviation white and must meet the applicable requirements of Sec. 
25.1397.
    (e) Light intensity. The minimum light intensities in all vertical 
planes, measured with the red filter (if used) and expressed in terms of 
``effective'' intensities, must meet the requirements of paragraph (f) 
of this section. The following relation must be assumed:
[GRAPHIC] [TIFF OMITTED] TC28SE91.049

where:

Ie=effective intensity (candles).
I(t)=instantaneous intensity as a function of time.
t2--t1=flash time interval (seconds).


Normally, the maximum value of effective intensity is obtained when t2 
and t1 are chosen so that the effective intensity is equal to the 
instantaneous intensity at t2 and t1.
    (f) Minimum effective intensities for anticollision lights. Each 
anticollision light effective intensity must equal or exceed the 
applicable values in the following table.

------------------------------------------------------------------------
                                                               Effective
          Angle above or below the horizontal plane            intensity
                                                               (candles)
------------------------------------------------------------------------
0[deg] to 5[deg]............................................         400
5[deg] to 10[deg]...........................................         240
10[deg] to 20[deg]..........................................          80
20[deg] to 30[deg]..........................................          40
30[deg] to 75[deg]..........................................          20
------------------------------------------------------------------------


[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-27, 
36 FR 12972, July 10, 1971; Amdt. 25-41, 42 FR 36970, July 18, 1977]



Sec. 25.1403  Wing icing detection lights.

    Unless operations at night in known or forecast icing conditions are 
prohibited by an operating limitation, a means must be provided for 
illuminating or otherwise determining the formation of ice on the parts 
of the wings that are critical from the standpoint of ice accumulation. 
Any illumination that is used must be of a type that will not cause 
glare or reflection that would handicap crewmembers in the performance 
of their duties.

[Amdt. 25-38, 41 FR 55468, Dec. 20, 1976]

                            Safety Equipment



Sec. 25.1411  General.

    (a) Accessibility. Required safety equipment to be used by the crew 
in an emergency must be readily accessible.
    (b) Stowage provisions. Stowage provisions for required emergency 
equipment must be furnished and must--
    (1) Be arranged so that the equipment is directly accessible and its 
location is obvious; and
    (2) Protect the safety equipment from inadvertent damage.
    (c) Emergency exit descent device. The stowage provisions for the 
emergency exit descent devices required by Sec. 25.810(a) must be at 
each exit for which they are intended.
    (d) Liferafts. (1) The stowage provisions for the liferafts 
described in Sec. 25.1415 must accommodate enough rafts for the maximum 
number of occupants for which certification for ditching is requested.
    (2) Liferafts must be stowed near exits through which the rafts can 
be launched during an unplanned ditching.
    (3) Rafts automatically or remotely released outside the airplane 
must be

[[Page 523]]

attached to the airplane by means of the static line prescribed in Sec. 
25.1415.
    (4) The stowage provisions for each portable liferaft must allow 
rapid detachment and removal of the raft for use at other than the 
intended exits.
    (e) Long-range signaling device. The stowage provisions for the 
long-range signaling device required by Sec. 25.1415 must be near an 
exit available during an unplanned ditching.
    (f) Life preserver stowage provisions. The stowage provisions for 
life preservers described in Sec. 25.1415 must accommodate one life 
preserver for each occupant for which certification for ditching is 
requested. Each life preserver must be within easy reach of each seated 
occupant.
    (g) Life line stowage provisions. If certification for ditching 
under Sec. 25.801 is requested, there must be provisions to store life 
lines. These provisions must--
    (1) Allow one life line to be attached to each side of the fuselage; 
and
    (2) Be arranged to allow the life lines to be used to enable the 
occupants to stay on the wing after ditching.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-32, 
37 FR 3972, Feb. 24, 1972; Amdt. 25-46, 43 FR 50598, Oct. 30, 1978; 
Amdt. 25-53, 45 FR 41593, June 19, 1980; Amdt. 25-70, 54 FR 43925, Oct. 
27, 1989; Amdt. 25-79, 58 FR 45229, Aug. 26, 1993; Amdt. 25-116, 69 FR 
62789, Oct. 27, 2004]



Sec. 25.1415  Ditching equipment.

    (a) Ditching equipment used in airplanes to be certificated for 
ditching under Sec. 25.801, and required by the operating rules of this 
chapter, must meet the requirements of this section.
    (b) Each liferaft and each life preserver must be approved. In 
addition--
    (1) Unless excess rafts of enough capacity are provided, the 
buoyancy and seating capacity beyond the rated capacity of the rafts 
must accommodate all occupants of the airplane in the event of a loss of 
one raft of the largest rated capacity; and
    (2) Each raft must have a trailing line, and must have a static line 
designed to hold the raft near the airplane but to release it if the 
airplane becomes totally submerged.
    (c) Approved survival equipment must be attached to each liferaft.
    (d) There must be an approved survival type emergency locator 
transmitter for use in one life raft.
    (e) For airplanes not certificated for ditching under Sec. 25.801 
and not having approved life preservers, there must be an approved 
flotation means for each occupant. This means must be within easy reach 
of each seated occupant and must be readily removable from the airplane.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-29, 
36 FR 18722, Sept. 21, 1971; Amdt. 25-50, 45 FR 38348, June 9, 1980; 
Amdt. 25-72, 55 FR 29785, July 20, 1990; Amdt. 25-82, 59 FR 32057, June 
21, 1994]



Sec. 25.1419  Ice protection.

    If the applicant seeks certification for flight in icing conditions, 
the airplane must be able to safely operate in the continuous maximum 
and intermittent maximum icing conditions of appendix C. To establish 
this--
    (a) An analysis must be performed to establish that the ice 
protection for the various components of the airplane is adequate, 
taking into account the various airplane operational configurations; and
    (b) To verify the ice protection analysis, to check for icing 
anomalies, and to demonstrate that the ice protection system and its 
components are effective, the airplane or its components must be flight 
tested in the various operational configurations, in measured natural 
atmospheric icing conditions and, as found necessary, by one or more of 
the following means:
    (1) Laboratory dry air or simulated icing tests, or a combination of 
both, of the components or models of the components.
    (2) Flight dry air tests of the ice protection system as a whole, or 
of its individual components.
    (3) Flight tests of the airplane or its components in measured 
simulated icing conditions.
    (c) Caution information, such as an amber caution light or 
equivalent, must be provided to alert the flightcrew when the anti-ice 
or de-ice system is not functioning normally.
    (d) For turbine engine powered airplanes, the ice protection 
provisions of this section are considered to be applicable primarily to 
the airframe. For

[[Page 524]]

the powerplant installation, certain additional provisions of subpart E 
of this part may be found applicable. (e) One of the following methods 
of icing detection and activation of the airframe ice protection system 
must be provided:
    (1) A primary ice detection system that automatically activates or 
alerts the flightcrew to activate the airframe ice protection system;
    (2) A definition of visual cues for recognition of the first sign of 
ice accretion on a specified surface combined with an advisory ice 
detection system that alerts the flightcrew to activate the airframe ice 
protection system; or
    (3) Identification of conditions conducive to airframe icing as 
defined by an appropriate static or total air temperature and visible 
moisture for use by the flightcrew to activate the airframe ice 
protection system.
    (f) Unless the applicant shows that the airframe ice protection 
system need not be operated during specific phases of flight, the 
requirements of paragraph (e) of this section are applicable to all 
phases of flight.
    (g) After the initial activation of the airframe ice protection 
system--
    (1) The ice protection system must be designed to operate 
continuously;
    (2) The airplane must be equipped with a system that automatically 
cycles the ice protection system; or
    (3) An ice detection system must be provided to alert the flightcrew 
each time the ice protection system must be cycled.
    (h) Procedures for operation of the ice protection system, including 
activation and deactivation, must be established and documented in the 
Airplane Flight Manual.

[Amdt. 25-72, 55 FR 29785, July 20, 1990, as amended by Amdt. 25-121, 72 
FR 44669, Aug. 8, 2007; Amdt. 25-129, 74 FR 38339, Aug. 3, 2009]



Sec. 25.1420  Supercooled large drop icing conditions.

    (a) If certification for flight in icing conditions is sought, in 
addition to the requirements of Sec. 25.1419, an airplane with a 
maximum takeoff weight less than 60,000 pounds or with reversible flight 
controls must be capable of operating in accordance with paragraphs 
(a)(1), (2), or (3), of this section.
    (1) Operating safely after encountering the icing conditions defined 
in Appendix O of this part:
    (i) The airplane must have a means to detect that it is operating in 
Appendix O icing conditions; and
    (ii) Following detection of Appendix O icing conditions, the 
airplane must be capable of operating safely while exiting all icing 
conditions.
    (2) Operating safely in a portion of the icing conditions defined in 
Appendix O of this part as selected by the applicant:
    (i) The airplane must have a means to detect that it is operating in 
conditions that exceed the selected portion of Appendix O icing 
conditions; and
    (ii) Following detection, the airplane must be capable of operating 
safely while exiting all icing conditions.
    (3) Operating safely in the icing conditions defined in Appendix O 
of this part.
    (b) To establish that the airplane can operate safely as required in 
paragraph (a) of this section, an applicant must show through analysis 
that the ice protection for the various components of the airplane is 
adequate, taking into account the various airplane operational 
configurations. To verify the analysis, one, or more as found necessary, 
of the following methods must be used:
    (1) Laboratory dry air or simulated icing tests, or a combination of 
both, of the components or models of the components.
    (2) Laboratory dry air or simulated icing tests, or a combination of 
both, of models of the airplane.
    (3) Flight tests of the airplane or its components in simulated 
icing conditions, measured as necessary to support the analysis.
    (4) Flight tests of the airplane with simulated ice shapes.
    (5) Flight tests of the airplane in natural icing conditions, 
measured as necessary to support the analysis.
    (c) For an airplane certified in accordance with paragraph (a)(2) or 
(3) of this section, the requirements of Sec. 25.1419(e), (f), (g), and 
(h) must be met for the icing conditions defined in Appendix O of this 
part in which the airplane is certified to operate.

[[Page 525]]

    (d) For the purposes of this section, the following definitions 
apply:
    (1) Reversible Flight Controls. Flight controls in the normal 
operating configuration that have force or motion originating at the 
airplane's control surface (for example, through aerodynamic loads, 
static imbalance, or trim or servo tab inputs) that is transmitted back 
to flight deck controls. This term refers to flight deck controls 
connected to the pitch, roll, or yaw control surfaces by direct 
mechanical linkages, cables, or push-pull rods in such a way that pilot 
effort produces motion or force about the hinge line.
    (2) Simulated Icing Test. Testing conducted in simulated icing 
conditions, such as in an icing tunnel or behind an icing tanker.
    (3) Simulated Ice Shape. Ice shape fabricated from wood, epoxy, or 
other materials by any construction technique.

[79 FR 65528, Nov. 4, 2014]

    Effective Date Note: At 79 FR 65528, Nov. 4, 2014, Sec. 25.1420 was 
added, effective Jan. 5, 2015.



Sec. 25.1421  Megaphones.

    If a megaphone is installed, a restraining means must be provided 
that is capable of restraining the megaphone when it is subjected to the 
ultimate inertia forces specified in Sec. 25.561(b)(3).

[Amdt. 25-41, 42 FR 36970, July 18, 1977]



Sec. 25.1423  Public address system.

    A public address system required by this chapter must--
    (a) Be powerable when the aircraft is in flight or stopped on the 
ground, after the shutdown or failure of all engines and auxiliary power 
units, or the disconnection or failure of all power sources dependent on 
their continued operation, for--
    (1) A time duration of at least 10 minutes, including an aggregate 
time duration of at least 5 minutes of announcements made by flight and 
cabin crewmembers, considering all other loads which may remain powered 
by the same source when all other power sources are inoperative; and
    (2) An additional time duration in its standby state appropriate or 
required for any other loads that are powered by the same source and 
that are essential to safety of flight or required during emergency 
conditions.
    (b) Be capable of operation within 3 seconds from the time a 
microphone is removed from its stowage.
    (c) Be intelligible at all passenger seats, lavatories, and flight 
attendant seats and work stations.
    (d) Be designed so that no unused, unstowed microphone will render 
the system inoperative.
    (e) Be capable of functioning independently of any required 
crewmember interphone system.
    (f) Be accessible for immediate use from each of two flight 
crewmember stations in the pilot compartment.
    (g) For each required floor-level passenger emergency exit which has 
an adjacent flight attendant seat, have a microphone which is readily 
accessible to the seated flight attendant, except that one microphone 
may serve more than one exit, provided the proximity of the exits allows 
unassisted verbal communication between seated flight attendants.

[Doc. No. 26003, 58 FR 45229, Aug. 26, 1993, as amended by Amdt. 25-115, 
69 FR 40527, July 2, 2004]

                         Miscellaneous Equipment



Sec. 25.1431  Electronic equipment.

    (a) In showing compliance with Sec. 25.1309 (a) and (b) with 
respect to radio and electronic equipment and their installations, 
critical environmental conditions must be considered.
    (b) Radio and electronic equipment must be supplied with power under 
the requirements of Sec. 25.1355(c).
    (c) Radio and electronic equipment, controls, and wiring must be 
installed so that operation of any one unit or system of units will not 
adversely affect the simultaneous operation of any other radio or 
electronic unit, or system of units, required by this chapter.
    (d) Electronic equipment must be designed and installed such that it 
does not cause essential loads to become inoperative as a result of 
electrical power supply transients or transients from other causes.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-113, 
69 FR 12530, Mar. 16, 2004]

[[Page 526]]



Sec. 25.1433  Vacuum systems.

    There must be means, in addition to the normal pressure relief, to 
automatically relieve the pressure in the discharge lines from the 
vacuum air pump when the delivery temperature of the air becomes unsafe.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29785, July 20, 1990]



Sec. 25.1435  Hydraulic systems.

    (a) Element design. Each element of the hydraulic system must be 
designed to:
    (1) Withstand the proof pressure without permanent deformation that 
would prevent it from performing its intended functions, and the 
ultimate pressure without rupture. The proof and ultimate pressures are 
defined in terms of the design operating pressure (DOP) as follows:

------------------------------------------------------------------------
                                                      Proof     Ultimate
                      Element                         (xDOP)     (xDOP)
------------------------------------------------------------------------
1. Tubes and fittings.............................        1.5        3.0
2. Pressure vessels containing gas:
  High pressure (e.g., accumulators)..............        3.0        4.0
  Low pressure (e.g., reservoirs).................        1.5        3.0
3. Hoses..........................................        2.0        4.0
4. All other elements.............................        1.5        2.0
------------------------------------------------------------------------

    (2) Withstand, without deformation that would prevent it from 
performing its intended function, the design operating pressure in 
combination with limit structural loads that may be imposed;
    (3) Withstand, without rupture, the design operating pressure 
multiplied by a factor of 1.5 in combination with ultimate structural 
load that can reasonably occur simultaneously;
    (4) Withstand the fatigue effects of all cyclic pressures, including 
transients, and associated externally induced loads, taking into account 
the consequences of element failure; and
    (5) Perform as intended under all environmental conditions for which 
the airplane is certificated.
    (b) System design. Each hydraulic system must:
    (1) Have means located at a flightcrew station to indicate 
appropriate system parameters, if
    (i) It performs a function necessary for continued safe flight and 
landing; or
    (ii) In the event of hydraulic system malfunction, corrective action 
by the crew to ensure continued safe flight and landing is necessary;
    (2) Have means to ensure that system pressures, including transient 
pressures and pressures from fluid volumetric changes in elements that 
are likely to remain closed long enough for such changes to occur, are 
within the design capabilities of each element, such that they meet the 
requirements defined in Sec. 25.1435(a)(1) through (a)(5);
    (3) Have means to minimize the release of harmful or hazardous 
concentrations of hydraulic fluid or vapors into the crew and passenger 
compartments during flight;
    (4) Meet the applicable requirements of Sec. Sec. 25.863, 25.1183, 
25.1185, and 25.1189 if a flammable hydraulic fluid is used; and
    (5) Be designed to use any suitable hydraulic fluid specified by the 
airplane manufacturer, which must be identified by appropriate markings 
as required by Sec. 25.1541.
    (c) Tests. Tests must be conducted on the hydraulic system(s), and/
or subsystem(s) and elements, except that analysis may be used in place 
of or to supplement testing, where the analysis is shown to be reliable 
and appropriate. All internal and external influences must be taken into 
account to an extent necessary to evaluate their effects, and to assure 
reliable system and element functioning and integration. Failure or 
unacceptable deficiency of an element or system must be corrected and be 
sufficiently retested, where necessary.
    (1) The system(s), subsystem(s), or element(s) must be subjected to 
performance, fatigue, and endurance tests representative of airplane 
ground and flight operations.
    (2) The complete system must be tested to determine proper 
functional performance and relation to the other systems, including 
simulation of relevant failure conditions, and to support or validate 
element design.
    (3) The complete hydraulic system(s) must be functionally tested on 
the airplane in normal operation over the

[[Page 527]]

range of motion of all associated user systems. The test must be 
conducted at the system relief pressure or 1.25 times the DOP if a 
system pressure relief device is not part of the system design. 
Clearances between hydraulic system elements and other systems or 
structural elements must remain adequate and there must be no 
detrimental effects.

[Doc. No. 28617, 66 FR 27402, May 16, 2001]



Sec. 25.1438  Pressurization and pneumatic systems.

    (a) Pressurization system elements must be burst pressure tested to 
2.0 times, and proof pressure tested to 1.5 times, the maximum normal 
operating pressure.
    (b) Pneumatic system elements must be burst pressure tested to 3.0 
times, and proof pressure tested to 1.5 times, the maximum normal 
operating pressure.
    (c) An analysis, or a combination of analysis and test, may be 
substituted for any test required by paragraph (a) or (b) of this 
section if the Administrator finds it equivalent to the required test.

[Amdt. 25-41, 42 FR 36971, July 18, 1977]



Sec. 25.1439  Protective breathing equipment.

    (a) Fixed (stationary, or built in) protective breathing equipment 
must be installed for the use of the flightcrew, and at least one 
portable protective breathing equipment shall be located at or near the 
flight deck for use by a flight crewmember. In addition, portable 
protective breathing equipment must be installed for the use of 
appropriate crewmembers for fighting fires in compartments accessible in 
flight other than the flight deck. This includes isolated compartments 
and upper and lower lobe galleys, in which crewmember occupancy is 
permitted during flight. Equipment must be installed for the maximum 
number of crewmembers expected to be in the area during any operation.
    (b) For protective breathing equipment required by paragraph (a) of 
this section or by the applicable Operating Regulations:
    (1) The equipment must be designed to protect the appropriate 
crewmember from smoke, carbon dioxide, and other harmful gases while on 
flight deck duty or while combating fires.
    (2) The equipment must include--
    (i) Masks covering the eyes, nose and mouth, or
    (ii) Masks covering the nose and mouth, plus accessory equipment to 
cover the eyes.
    (3) Equipment, including portable equipment, must allow 
communication with other crewmembers while in use. Equipment available 
at flightcrew assigned duty stations must also enable the flightcrew to 
use radio equipment.
    (4) The part of the equipment protecting the eyes shall not cause 
any appreciable adverse effect on vision and must allow corrective 
glasses to be worn.
    (5) The equipment must supply protective oxygen of 15 minutes 
duration per crewmember at a pressure altitude of 8,000 feet with a 
respiratory minute volume of 30 liters per minute BTPD. The equipment 
and system must be designed to prevent any inward leakage to the inside 
of the device and prevent any outward leakage causing significant 
increase in the oxygen content of the local ambient atmosphere. If a 
demand oxygen system is used, a supply of 300 liters of free oxygen at 
70 [deg]F. and 760 mm. Hg. pressure is considered to be of 15-minute 
duration at the prescribed altitude and minute volume. If a continuous 
flow open circuit protective breathing system is used, a flow rate of 60 
liters per minute at 8,000 feet (45 liters per minute at sea level) and 
a supply of 600 liters of free oxygen at 70 [deg]F. and 760 mm. Hg. 
pressure is considered to be of 15-minute duration at the prescribed 
altitude and minute volume. Continuous flow systems must not increase 
the ambient oxygen content of the local atmosphere above that of demand 
systems. BTPD refers to body temperature conditions (that is, 37 
[deg]C., at ambient pressure, dry).
    (6) The equipment must meet the requirements of Sec. 25.1441.

[Doc. No. FAA-2002-13859, 69 FR 40528, July 2, 2004]

[[Page 528]]



Sec. 25.1441  Oxygen equipment and supply.

    (a) If certification with supplemental oxygen equipment is 
requested, the equipment must meet the requirements of this section and 
Sec. Sec. 25.1443 through 25.1453.
    (b) The oxygen system must be free from hazards in itself, in its 
method of operation, and in its effect upon other components.
    (c) There must be a means to allow the crew to readily determine, 
during flight, the quantity of oxygen available in each source of 
supply.
    (d) The oxygen flow rate and the oxygen equipment for airplanes for 
which certification for operation above 40,000 feet is requested must be 
approved.



Sec. 25.1443  Minimum mass flow of supplemental oxygen.

    (a) If continuous flow equipment is installed for use by flight 
crewmembers, the minimum mass flow of supplemental oxygen required for 
each crewmember may not be less than the flow required to maintain, 
during inspiration, a mean tracheal oxygen partial pressure of 149 mm. 
Hg. when breathing 15 liters per minute, BTPS, and with a maximum tidal 
volume of 700 cc. with a constant time interval between respirations.
    (b) If demand equipment is installed for use by flight crewmembers, 
the minimum mass flow of supplemental oxygen required for each 
crewmember may not be less than the flow required to maintain, during 
inspiration, a mean tracheal oxygen partial pressure of 122 mm. Hg., up 
to and including a cabin pressure altitude of 35,000 feet, and 95 
percent oxygen between cabin pressure altitudes of 35,000 and 40,000 
feet, when breathing 20 liters per minute BTPS. In addition, there must 
be means to allow the crew to use undiluted oxygen at their discretion.
    (c) For passengers and cabin attendants, the minimum mass flow of 
supplemental oxygen required for each person at various cabin pressure 
altitudes may not be less than the flow required to maintain, during 
inspiration and while using the oxygen equipment (including masks) 
provided, the following mean tracheal oxygen partial pressures:
    (1) At cabin pressure altitudes above 10,000 feet up to and 
including 18,500 feet, a mean tracheal oxygen partial pressure of 100 
mm. Hg. when breathing 15 liters per minute, BTPS, and with a tidal 
volume of 700 cc. with a constant time interval between respirations.
    (2) At cabin pressure altitudes above 18,500 feet up to and 
including 40,000 feet, a mean tracheal oxygen partial pressure of 83.8 
mm. Hg. when breathing 30 liters per minute, BTPS, and with a tidal 
volume of 1,100 cc. with a constant time interval between respirations.
    (d) If first-aid oxygen equipment is installed, the minimum mass 
flow of oxygen to each user may not be less than four liters per minute, 
STPD. However, there may be a means to decrease this flow to not less 
than two liters per minute, STPD, at any cabin altitude. The quantity of 
oxygen required is based upon an average flow rate of three liters per 
minute per person for whom first-aid oxygen is required.
    (e) If portable oxygen equipment is installed for use by 
crewmembers, the minimum mass flow of supplemental oxygen is the same as 
specified in paragraph (a) or (b) of this section, whichever is 
applicable.



Sec. 25.1445  Equipment standards for the oxygen distributing system.

    (a) When oxygen is supplied to both crew and passengers, the 
distribution system must be designed for either--
    (1) A source of supply for the flight crew on duty and a separate 
source for the passengers and other crewmembers; or
    (2) A common source of supply with means to separately reserve the 
minimum supply required by the flight crew on duty.
    (b) Portable walk-around oxygen units of the continuous flow, 
diluter-demand, and straight demand kinds may be used to meet the crew 
or passenger breathing requirements.



Sec. 25.1447  Equipment standards for oxygen dispensing units.

    If oxygen dispensing units are installed, the following apply:
    (a) There must be an individual dispensing unit for each occupant 
for

[[Page 529]]

whom supplemental oxygen is to be supplied. Units must be designed to 
cover the nose and mouth and must be equipped with a suitable means to 
retain the unit in position on the face. Flight crew masks for 
supplemental oxygen must have provisions for the use of communication 
equipment.
    (b) If certification for operation up to and including 25,000 feet 
is requested, an oxygen supply terminal and unit of oxygen dispensing 
equipment for the immediate use of oxygen by each crewmember must be 
within easy reach of that crewmember. For any other occupants, the 
supply terminals and dispensing equipment must be located to allow the 
use of oxygen as required by the operating rules in this chapter.
    (c) If certification for operation above 25,000 feet is requested, 
there must be oxygen dispensing equipment meeting the following 
requirements:
    (1) There must be an oxygen dispensing unit connected to oxygen 
supply terminals immediately available to each occupant, wherever 
seated, and at least two oxygen dispensing units connected to oxygen 
terminals in each lavatory. The total number of dispensing units and 
outlets in the cabin must exceed the number of seats by at least 10 
percent. The extra units must be as uniformly distributed throughout the 
cabin as practicable. If certification for operation above 30,000 feet 
is requested, the dispensing units providing the required oxygen flow 
must be automatically presented to the occupants before the cabin 
pressure altitude exceeds 15,000 feet. The crew must be provided with a 
manual means of making the dispensing units immediately available in the 
event of failure of the automatic system.
    (2) Each flight crewmember on flight deck duty must be provided with 
a quick-donning type oxygen dispensing unit connected to an oxygen 
supply terminal. This dispensing unit must be immediately available to 
the flight crewmember when seated at his station, and installed so that 
it:
    (i) Can be placed on the face from its ready position, properly 
secured, sealed, and supplying oxygen upon demand, with one hand, within 
five seconds and without disturbing eyeglasses or causing delay in 
proceeding with emergency duties; and
    (ii) Allows, while in place, the performance of normal communication 
functions.
    (3) The oxygen dispensing equipment for the flight crewmembers must 
be:
    (i) The diluter demand or pressure demand (pressure demand mask with 
a diluter demand pressure breathing regulator) type, or other approved 
oxygen equipment shown to provide the same degree of protection, for 
airplanes to be operated above 25,000 feet.
    (ii) The pressure demand (pressure demand mask with a diluter demand 
pressure breathing regulator) type with mask-mounted regulator, or other 
approved oxygen equipment shown to provide the same degree of 
protection, for airplanes operated at altitudes where decompressions 
that are not extremely improbable may expose the flightcrew to cabin 
pressure altitudes in excess of 34,000 feet.
    (4) Portable oxygen equipment must be immediately available for each 
cabin attendant. The portable oxygen equipment must have the oxygen 
dispensing unit connected to the portable oxygen supply.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 
42 FR 36971, July 18, 1977; Amdt. 25-87, 61 FR 28696, June 5, 1996; 
Amdt. 25-116, 69 FR 62789, Oct. 27, 2004]



Sec. 25.1449  Means for determining use of oxygen.

    There must be a means to allow the crew to determine whether oxygen 
is being delivered to the dispensing equipment.



Sec. 25.1450  Chemical oxygen generators.

    (a) For the purpose of this section, a chemical oxygen generator is 
defined as a device which produces oxygen by chemical reaction.
    (b) Each chemical oxygen generator must be designed and installed in 
accordance with the following requirements:
    (1) Surface temperature developed by the generator during operation 
may not create a hazard to the airplane or to its occupants.
    (2) Means must be provided to relieve any internal pressure that may 
be hazardous.

[[Page 530]]

    (3) Except as provided in SFAR 109, each chemical oxygen generator 
installation must meet the requirements of Sec. 25.795(d).
    (c) In addition to meeting the requirements in paragraph (b) of this 
section, each portable chemical oxygen generator that is capable of 
sustained operation by successive replacement of a generator element 
must be placarded to show--
    (1) The rate of oxygen flow, in liters per minute;
    (2) The duration of oxygen flow, in minutes, for the replaceable 
generator element; and
    (3) A warning that the replaceable generator element may be hot, 
unless the element construction is such that the surface temperature 
cannot exceed 100 degrees F.

[Amdt. 25-41, 42 FR 36971, July 18, 1977, as amended at 79 FR 13519, 
Mar. 11, 2014]



Sec. 25.1453  Protection of oxygen equipment from rupture.

    Oxygen pressure tanks, and lines between tanks and the shutoff 
means, must be--
    (a) Protected from unsafe temperatures; and
    (b) Located where the probability and hazards of rupture in a crash 
landing are minimized.



Sec. 25.1455  Draining of fluids subject to freezing.

    If fluids subject to freezing may be drained overboard in flight or 
during ground operation, the drains must be designed and located to 
prevent the formation of hazardous quantities of ice on the airplane as 
a result of the drainage.

[Amdt. 25-23, 35 FR 5680, Apr. 8, 1970]



Sec. 25.1457  Cockpit voice recorders.

    (a) Each cockpit voice recorder required by the operating rules of 
this chapter must be approved and must be installed so that it will 
record the following:
    (1) Voice communications transmitted from or received in the 
airplane by radio.
    (2) Voice communications of flight crewmembers on the flight deck.
    (3) Voice communications of flight crewmembers on the flight deck, 
using the airplane's interphone system.
    (4) Voice or audio signals identifying navigation or approach aids 
introduced into a headset or speaker.
    (5) Voice communications of flight crewmembers using the passenger 
loudspeaker system, if there is such a system and if the fourth channel 
is available in accordance with the requirements of paragraph (c)(4)(ii) 
of this section.
    (6) If datalink communication equipment is installed, all datalink 
communications, using an approved data message set. Datalink messages 
must be recorded as the output signal from the communications unit that 
translates the signal into usable data.
    (b) The recording requirements of paragraph (a)(2) of this section 
must be met by installing a cockpit-mounted area microphone, located in 
the best position for recording voice communications originating at the 
first and second pilot stations and voice communications of other 
crewmembers on the flight deck when directed to those stations. The 
microphone must be so located and, if necessary, the preamplifiers and 
filters of the recorder must be so adjusted or supplemented, that the 
intelligibility of the recorded communications is as high as practicable 
when recorded under flight cockpit noise conditions and played back. 
Repeated aural or visual playback of the record may be used in 
evaluating intelligibility.
    (c) Each cockpit voice recorder must be installed so that the part 
of the communication or audio signals specified in paragraph (a) of this 
section obtained from each of the following sources is recorded on a 
separate channel:
    (1) For the first channel, from each boom, mask, or hand-held 
microphone, headset, or speaker used at the first pilot station.
    (2) For the second channel from each boom, mask, or hand-held 
microphone, headset, or speaker used at the second pilot station.
    (3) For the third channel--from the cockpit-mounted area microphone.
    (4) For the fourth channel, from--

[[Page 531]]

    (i) Each boom, mask, or hand-held microphone, headset, or speaker 
used at the station for the third and fourth crew members; or
    (ii) If the stations specified in paragraph (c)(4)(i) of this 
section are not required or if the signal at such a station is picked up 
by another channel, each microphone on the flight deck that is used with 
the passenger loudspeaker system, if its signals are not picked up by 
another channel.
    (5) As far as is practicable all sounds received by the microphone 
listed in paragraphs (c)(1), (2), and (4) of this section must be 
recorded without interruption irrespective of the position of the 
interphone-transmitter key switch. The design shall ensure that sidetone 
for the flight crew is produced only when the interphone, public address 
system, or radio transmitters are in use.
    (d) Each cockpit voice recorder must be installed so that--
    (1)(i) It receives its electrical power from the bus that provides 
the maximum reliability for operation of the cockpit voice recorder 
without jeopardizing service to essential or emergency loads.
    (ii) It remains powered for as long as possible without jeopardizing 
emergency operation of the airplane.
    (2) There is an automatic means to simultaneously stop the recorder 
and prevent each erasure feature from functioning, within 10 minutes 
after crash impact;
    (3) There is an aural or visual means for preflight checking of the 
recorder for proper operation;
    (4) Any single electrical failure external to the recorder does not 
disable both the cockpit voice recorder and the flight data recorder;
    (5) It has an independent power source--
    (i) That provides 10 1 minutes of electrical 
power to operate both the cockpit voice recorder and cockpit-mounted 
area microphone;
    (ii) That is located as close as practicable to the cockpit voice 
recorder; and
    (iii) To which the cockpit voice recorder and cockpit-mounted area 
microphone are switched automatically in the event that all other power 
to the cockpit voice recorder is interrupted either by normal shutdown 
or by any other loss of power to the electrical power bus; and
    (6) It is in a separate container from the flight data recorder when 
both are required. If used to comply with only the cockpit voice 
recorder requirements, a combination unit may be installed.
    (e) The recorder container must be located and mounted to minimize 
the probability of rupture of the container as a result of crash impact 
and consequent heat damage to the recorder from fire.
    (1) Except as provided in paragraph (e)(2) of this section, the 
recorder container must be located as far aft as practicable, but need 
not be outside of the pressurized compartment, and may not be located 
where aft-mounted engines may crush the container during impact.
    (2) If two separate combination digital flight data recorder and 
cockpit voice recorder units are installed instead of one cockpit voice 
recorder and one digital flight data recorder, the combination unit that 
is installed to comply with the cockpit voice recorder requirements may 
be located near the cockpit.
    (f) If the cockpit voice recorder has a bulk erasure device, the 
installation must be designed to minimize the probability of inadvertent 
operation and actuation of the device during crash impact.
    (g) Each recorder container must--
    (1) Be either bright orange or bright yellow;
    (2) Have reflective tape affixed to its external surface to 
facilitate its location under water; and
    (3) Have an underwater locating device, when required by the 
operating rules of this chapter, on or adjacent to the container which 
is secured in such manner that they are not likely to be separated 
during crash impact.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-2, 30 
FR 3932, Mar. 26, 1965; Amdt. 25-16, 32 FR 13914, Oct. 6, 1967; Amdt. 
25-41, 42 FR 36971, July 18, 1977; Amdt. 25-65, 53 FR 26143, July 11, 
1988; Amdt. 25-124, 73 FR 12563, Mar. 7, 2008; 74 FR 32800, July 9, 
2009]

[[Page 532]]



Sec. 25.1459  Flight data recorders.

    (a) Each flight recorder required by the operating rules of this 
chapter must be installed so that--
    (1) It is supplied with airspeed, altitude, and directional data 
obtained from sources that meet the accuracy requirements of Sec. Sec. 
25.1323, 25.1325, and 25.1327, as appropriate;
    (2) The vertical acceleration sensor is rigidly attached, and 
located longitudinally either within the approved center of gravity 
limits of the airplane, or at a distance forward or aft of these limits 
that does not exceed 25 percent of the airplane's mean aerodynamic 
chord;
    (3)(i) It receives its electrical power from the bus that provides 
the maximum reliability for operation of the flight data recorder 
without jeopardizing service to essential or emergency loads.
    (ii) It remains powered for as long as possible without jeopardizing 
emergency operation of the airplane.
    (4) There is an aural or visual means for preflight checking of the 
recorder for proper recording of data in the storage medium;
    (5) Except for recorders powered solely by the engine-driven 
electrical generator system, there is an automatic means to 
simultaneously stop a recorder that has a data erasure feature and 
prevent each erasure feature from functioning, within 10 minutes after 
crash impact;
    (6) There is a means to record data from which the time of each 
radio transmission either to or from ATC can be determined;
    (7) Any single electrical failure external to the recorder does not 
disable both the cockpit voice recorder and the flight data recorder; 
and
    (8) It is in a separate container from the cockpit voice recorder 
when both are required. If used to comply with only the flight data 
recorder requirements, a combination unit may be installed. If a 
combination unit is installed as a cockpit voice recorder to comply with 
Sec. 25.1457(e)(2), a combination unit must be used to comply with this 
flight data recorder requirement.
    (b) Each nonejectable record container must be located and mounted 
so as to minimize the probability of container rupture resulting from 
crash impact and subsequent damage to the record from fire. In meeting 
this requirement the record container must be located as far aft as 
practicable, but need not be aft of the pressurized compartment, and may 
not be where aft-mounted engines may crush the container upon impact.
    (c) A correlation must be established between the flight recorder 
readings of airspeed, altitude, and heading and the corresponding 
readings (taking into account correction factors) of the first pilot's 
instruments. The correlation must cover the airspeed range over which 
the airplane is to be operated, the range of altitude to which the 
airplane is limited, and 360 degrees of heading. Correlation may be 
established on the ground as appropriate.
    (d) Each recorder container must--
    (1) Be either bright orange or bright yellow;
    (2) Have reflective tape affixed to its external surface to 
facilitate its location under water; and
    (3) Have an underwater locating device, when required by the 
operating rules of this chapter, on or adjacent to the container which 
is secured in such a manner that they are not likely to be separated 
during crash impact.
    (e) Any novel or unique design or operational characteristics of the 
aircraft shall be evaluated to determine if any dedicated parameters 
must be recorded on flight recorders in addition to or in place of 
existing requirements.

[Amdt. 25-8, 31 FR 127, Jan. 6, 1966, as amended by Amdt. 25-25, 35 FR 
13192, Aug. 19, 1970; Amdt. 25-37, 40 FR 2577, Jan. 14, 1975; Amdt. 25-
41, 42 FR 36971, July 18, 1977; Amdt. 25-65, 53 FR 26144, July 11, 1988; 
Amdt. 25-124, 73 FR 12563, Mar. 7, 2008; 74 FR 32800, July 9, 2009]



Sec. 25.1461  Equipment containing high energy rotors.

    (a) Equipment containing high energy rotors must meet paragraph (b), 
(c), or (d) of this section.
    (b) High energy rotors contained in equipment must be able to 
withstand damage caused by malfunctions, vibration, abnormal speeds, and 
abnormal temperatures. In addition--
    (1) Auxiliary rotor cases must be able to contain damage caused by 
the failure of high energy rotor blades; and

[[Page 533]]

    (2) Equipment control devices, systems, and instrumentation must 
reasonably ensure that no operating limitations affecting the integrity 
of high energy rotors will be exceeded in service.
    (c) It must be shown by test that equipment containing high energy 
rotors can contain any failure of a high energy rotor that occurs at the 
highest speed obtainable with the normal speed control devices 
inoperative.
    (d) Equipment containing high energy rotors must be located where 
rotor failure will neither endanger the occupants nor adversely affect 
continued safe flight.

[Amdt. 25-41, 42 FR 36971, July 18, 1977]



             Subpart G_Operating Limitations and Information



Sec. 25.1501  General.

    (a) Each operating limitation specified in Sec. Sec. 25.1503 
through 25.1533 and other limitations and information necessary for safe 
operation must be established.
    (b) The operating limitations and other information necessary for 
safe operation must be made available to the crewmembers as prescribed 
in Sec. Sec. 25.1541 through 25.1587.

[Amdt. 25-42, 43 FR 2323, Jan. 16, 1978]

                          Operating Limitations



Sec. 25.1503  Airspeed limitations: general.

    When airspeed limitations are a function of weight, weight 
distribution, altitude, or Mach number, limitations corresponding to 
each critical combination of these factors must be established.



Sec. 25.1505  Maximum operating limit speed.

    The maximum operating limit speed (VMO/MMO airspeed or Mach Number, 
whichever is critical at a particular altitude) is a speed that may not 
be deliberately exceeded in any regime of flight (climb, cruise, or 
descent), unless a higher speed is authorized for flight test or pilot 
training operations. VMO/MMO must be established so that it is not 
greater than the design cruising speed VC and so that it is sufficiently 
below VD/MD or VDF/MDF, to make it highly improbable that the latter 
speeds will be inadvertently exceeded in operations. The speed margin 
between VMO/MMO and VD/MD or VDFM/DF may not be less than that 
determined under Sec. 25.335(b) or found necessary during the flight 
tests conducted under Sec. 25.253.

[Amdt. 25-23, 35 FR 5680, Apr. 8, 1970]



Sec. 25.1507  Maneuvering speed.

    The maneuvering speed must be established so that it does not exceed 
the design maneuvering speed VA determined under Sec. 25.335(c).



Sec. 25.1511  Flap extended speed.

    The established flap extended speed VFE must be established so that 
it does not exceed the design flap speed VF chosen under Sec. Sec. 
25.335(e) and 25.345, for the corresponding flap positions and engine 
powers.



Sec. 25.1513  Minimum control speed.

    The minimum control speed VMC determined under Sec. 25.149 must be 
established as an operating limitation.



Sec. 25.1515  Landing gear speeds.

    (a) The established landing gear operating speed or speeds, VLO, may 
not exceed the speed at which it is safe both to extend and to retract 
the landing gear, as determined under Sec. 25.729 or by flight 
characteristics. If the extension speed is not the same as the 
retraction speed, the two speeds must be designated as VLO(EXT) and 
VLO(RET), respectively.
    (b) The established landing gear extended speed VLE may not exceed 
the speed at which it is safe to fly with the landing gear secured in 
the fully extended position, and that determined under Sec. 25.729.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55468, Dec. 20, 1976]



Sec. 25.1516  Other speed limitations.

    Any other limitation associated with speed must be established.

[Doc. No. 2000-8511, 66 FR 34024, June 26, 2001]

[[Page 534]]



Sec. 25.1517  Rough air speed, VRA.

    A rough air speed, VRA, for use as the recommended 
turbulence penetration airspeed in Sec. 25.1585(a)(8), must be 
established, which--
    (1) Is not greater than the design airspeed for maximum gust 
intensity, selected for VB; and
    (2) Is not less than the minimum value of VB specified in 
Sec. 25.335(d); and
    (3) Is sufficiently less than VMO to ensure that likely 
speed variation during rough air encounters will not cause the overspeed 
warning to operate too frequently. In the absence of a rational 
investigation substantiating the use of other values, VRA 
must be less than VMO--35 knots (TAS).

[Doc. No. 27902, 61 FR 5222, Feb. 9, 1996]

    Effective Date Note: At 79 FR 73469, Dec. 11, 2014, Sec. 25.1517 
was revised, effective Feb. 9, 2015. For the convenience of the user, 
the revised text is set forth as follows:



Sec. 25.1517  Rough air speed, VRA.

    (a) A rough air speed, VRA, for use as the recommended 
turbulence penetration airspeed, and a rough air Mach number, 
MRA, for use as the recommended turbulence penetration Mach 
number, must be established. VRA/MRA must be 
sufficiently less than VMO/MMO to ensure that 
likely speed variation during rough air encounters will not cause the 
overspeed warning to operate too frequently.
    (b) At altitudes where VMO is not limited by Mach number, 
in the absence of a rational investigation substantiating the use of 
other values, VRA must be less than VMO--35 KTAS.
    (c) At altitudes where VMO is limited by Mach number, 
MRA may be chosen to provide an optimum margin between low 
and high speed buffet boundaries.



Sec. 25.1519  Weight, center of gravity, and weight distribution.

    The airplane weight, center of gravity, and weight distribution 
limitations determined under Sec. Sec. 25.23 through 25.27 must be 
established as operating limitations.



Sec. 25.1521  Powerplant limitations.

    (a) General. The powerplant limitations prescribed in this section 
must be established so that they do not exceed the corresponding limits 
for which the engines or propellers are type certificated and do not 
exceed the values on which compliance with any other requirement of this 
part is based.
    (b) Reciprocating engine installations. Operating limitations 
relating to the following must be established for reciprocating engine 
installations:
    (1) Horsepower or torque, r.p.m., manifold pressure, and time at 
critical pressure altitude and sea level pressure altitude for--
    (i) Maximum continuous power (relating to unsupercharged operation 
or to operation in each supercharger mode as applicable); and
    (ii) Takeoff power (relating to unsupercharged operation or to 
operation in each supercharger mode as applicable).
    (2) Fuel grade or specification.
    (3) Cylinder head and oil temperatures.
    (4) Any other parameter for which a limitation has been established 
as part of the engine type certificate except that a limitation need not 
be established for a parameter that cannot be exceeded during normal 
operation due to the design of the installation or to another 
established limitation.
    (c) Turbine engine installations. Operating limitations relating to 
the following must be established for turbine engine installations:
    (1) Horsepower, torque or thrust, r.p.m., gas temperature, and time 
for--
    (i) Maximum continuous power or thrust (relating to augmented or 
unaugmented operation as applicable).
    (ii) Takeoff power or thrust (relating to augmented or unaugmented 
operation as applicable).
    (2) Fuel designation or specification.
    (3) Any other parameter for which a limitation has been established 
as part of the engine type certificate except that a limitation need not 
be established for a parameter that cannot be exceeded during normal 
operation due to the design of the installation or to another 
established limitation.
    (d) Ambient temperature. An ambient temperature limitation 
(including limitations for winterization installations, if applicable) 
must be established as the maximum ambient atmospheric temperature 
established in accordance with Sec. 25.1043(b).

[Amdt. 25-72, 55 FR 29786, July 20, 1990]

    Effective Date Note: At 79 FR 65528, Nov. 4, 2014, Sec. 25.1521 was 
amended by redesignating paragraph (c)(3) as paragraph (c)(4), revising 
newly redesignated paragraph (c)(4),

[[Page 535]]

and adding new paragraph (c)(3), effective Jan. 5, 2015. For the 
convenience of the user, the added and revised text is set forth as 
follows:



Sec. 25.1521  Powerplant limitations.

                                * * * * *

    (c) * * *
    (3) Maximum time interval between engine run-ups from idle, run-up 
power setting and duration at power for ground operation in icing 
conditions, as defined in Sec. 25.1093(b)(2).
    (4) Any other parameter for which a limitation has been established 
as part of the engine type certificate except that a limitation need not 
be established for a parameter that cannot be exceeded during normal 
operation due to the design of the installation or to another 
established limitation.

                                * * * * *



Sec. 25.1522  Auxiliary power unit limitations.

    If an auxiliary power unit is installed in the airplane, limitations 
established for the auxiliary power unit, including categories of 
operation, must be specified as operating limitations for the airplane.

[Amdt. 25-72, 55 FR 29786, July 20, 1990]



Sec. 25.1523  Minimum flight crew.

    The minimum flight crew must be established so that it is sufficient 
for safe operation, considering--
    (a) The workload on individual crewmembers;
    (b) The accessibility and ease of operation of necessary controls by 
the appropriate crewmember; and
    (c) The kind of operation authorized under Sec. 25.1525.

The criteria used in making the determinations required by this section 
are set forth in appendix D.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-3, 30 
FR 6067, Apr. 29, 1965]



Sec. 25.1525  Kinds of operation.

    The kinds of operation to which the airplane is limited are 
established by the category in which it is eligible for certification 
and by the installed equipment.



Sec. 25.1527  Ambient air temperature and operating altitude.

    The extremes of the ambient air temperature and operating altitude 
for which operation is allowed, as limited by flight, structural, 
powerplant, functional, or equipment characteristics, must be 
established.

[Doc. No. 2000-8511, 66 FR 34024, June 26, 2001]



Sec. 25.1529  Instructions for Continued Airworthiness.

    The applicant must prepare Instructions for Continued Airworthiness 
in accordance with appendix H to this part that are acceptable to the 
Administrator. The instructions may be incomplete at type certification 
if a program exists to ensure their completion prior to delivery of the 
first airplane or issuance of a standard certificate of airworthiness, 
whichever occurs later.

[Amdt. 25-54, 45 FR 60173, Sept. 11, 1980]



Sec. 25.1531  Maneuvering flight load factors.

    Load factor limitations, not exceeding the positive limit load 
factors determined from the maneuvering diagram in Sec. 25.333(b), must 
be established.



Sec. 25.1533  Additional operating limitations.

    (a) Additional operating limitations must be established as follows:
    (1) The maximum takeoff weights must be established as the weights 
at which compliance is shown with the applicable provisions of this part 
(including the takeoff climb provisions of Sec. 25.121(a) through (c), 
for altitudes and ambient temperatures).
    (2) The maximum landing weights must be established as the weights 
at which compliance is shown with the applicable provisions of this part 
(including the landing and approach climb provisions of Sec. Sec. 
25.119 and 25.121(d) for altitudes and ambient temperatures).
    (3) The minimum takeoff distances must be established as the 
distances at which compliance is shown with the applicable provisions of 
this part (including the provisions of Sec. Sec. 25.109 and

[[Page 536]]

25.113, for weights, altitudes, temperatures, wind components, runway 
surface conditions (dry and wet), and runway gradients) for smooth, 
hard-surfaced runways. Additionally, at the option of the applicant, wet 
runway takeoff distances may be established for runway surfaces that 
have been grooved or treated with a porous friction course, and may be 
approved for use on runways where such surfaces have been designed 
constructed, and maintained in a manner acceptable to the Administrator.
    (b) The extremes for variable factors (such as altitude, 
temperature, wind, and runway gradients) are those at which compliance 
with the applicable provisions of this part is shown.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 
41 FR 55468, Dec. 20, 1976; Amdt. 25-72, 55 FR 29786, July 20, 1990; 
Amdt. 25-92, 63 FR 8321, Feb. 18, 1998]

    Effective Date Note: At 79 FR 65528, Nov. 4, 2014, Sec. 25.1533 was 
amended by adding a new paragraph (c), effective Jan. 5, 2015. For the 
convenience of the user, the added text is set forth as follows:



Sec. 25.1533  Additional operating limitations.

                                * * * * *

    (c) For airplanes certified in accordance with Sec. 25.1420(a)(1) 
or (2), an operating limitation must be established to:
    (1) Prohibit intentional flight, including takeoff and landing, into 
icing conditions defined in Appendix O of this part for which the 
airplane has not been certified to safely operate; and
    (2) Require exiting all icing conditions if icing conditions defined 
in Appendix O of this part are encountered for which the airplane has 
not been certified to safely operate.



Sec. 25.1535  ETOPS approval.

    Except as provided in Sec. 25.3, each applicant seeking ETOPS type 
design approval must comply with the provisions of Appendix K of this 
part.

[Doc. No. FAA-2002-6717, 72 FR 1873, Jan. 16, 2007]

                          Markings and Placards



Sec. 25.1541  General.

    (a) The airplane must contain--
    (1) The specified markings and placards; and
    (2) Any additional information, instrument markings, and placards 
required for the safe operation if there are unusual design, operating, 
or handling characteristics.
    (b) Each marking and placard prescribed in paragraph (a) of this 
section--
    (1) Must be displayed in a conspicuous place; and
    (2) May not be easily erased, disfigured, or obscured.



Sec. 25.1543  Instrument markings: general.

    For each instrument--
    (a) When markings are on the cover glass of the instrument, there 
must be means to maintain the correct alignment of the glass cover with 
the face of the dial; and
    (b) Each instrument marking must be clearly visible to the 
appropriate crewmember.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 
55 FR 29786, July 20, 1990]



Sec. 25.1545  Airspeed limitation information.

    The airspeed limitations required by Sec. 25.1583 (a) must be 
easily read and understood by the flight crew.



Sec. 25.1547  Magnetic direction indicator.

    (a) A placard meeting the requirements of this section must be 
installed on, or near, the magnetic direction indicator.
    (b) The placard must show the calibration of the instrument in level 
flight with the engines operating.
    (c) The placard must state whether the calibration was made with 
radio receivers on or off.
    (d) Each calibration reading must be in terms of magnetic heading in 
not more than 45 degree increments.



Sec. 25.1549  Powerplant and auxiliary power unit instruments.

    For each required powerplant and auxiliary power unit instrument, as 
appropriate to the type of instrument--
    (a) Each maximum and, if applicable, minimum safe operating limit 
must be marked with a red radial or a red line;
    (b) Each normal operating range must be marked with a green arc or

[[Page 537]]

green line, not extending beyond the maximum and minimum safe limits;
    (c) Each takeoff and precautionary range must be marked with a 
yellow arc or a yellow line; and
    (d) Each engine, auxiliary power unit, or propeller speed range that 
is restricted because of excessive vibration stresses must be marked 
with red arcs or red lines.

[Amdt. 25-40, 42 FR 15044, Mar. 17, 1977]



Sec. 25.1551  Oil quantity indication.

    Each oil quantity indicating means must be marked to indicate the 
quantity of oil readily and accurately.

[Amdt. 25-72, 55 FR 29786, July 20, 1990]



Sec. 25.1553  Fuel quantity indicator.

    If the unusable fuel supply for any tank exceeds one gallon, or five 
percent of the tank capacity, whichever is greater, a red arc must be 
marked on its indicator extending from the calibrated zero reading to 
the lowest reading obtainable in level flight.



Sec. 25.1555  Control markings.

    (a) Each cockpit control, other than primary flight controls and 
controls whose function is obvious, must be plainly marked as to its 
function and method of operation.
    (b) Each aerodynamic control must be marked under the requirements 
of Sec. Sec. 25.677 and 25.699.
    (c) For powerplant fuel controls--
    (1) Each fuel tank selector control must be marked to indicate the 
position corresponding to each tank and to each existing cross feed 
position;
    (2) If safe operation requires the use of any tanks in a specific 
sequence, that sequence must be marked on, or adjacent to, the selector 
for those tanks; and
    (3) Each valve control for each engine must be marked to indicate 
the position corresponding to each engine controlled.
    (d) For accessory, auxiliary, and emergency controls--
    (1) Each emergency control (including each fuel jettisoning and 
fluid shutoff must be colored red; and
    (2) Each visual indicator required by Sec. 25.729(e) must be marked 
so that the pilot can determine at any time when the wheels are locked 
in either extreme position, if retractable landing gear is used.



Sec. 25.1557  Miscellaneous markings and placards.

    (a) Baggage and cargo compartments and ballast location. Each 
baggage and cargo compartment, and each ballast location must have a 
placard stating any limitations on contents, including weight, that are 
necessary under the loading requirements. However, underseat 
compartments designed for the storage of carry-on articles weighing not 
more than 20 pounds need not have a loading limitation placard.
    (b) Powerplant fluid filler openings. The following apply:
    (1) Fuel filler openings must be marked at or near the filler cover 
with--
    (i) The word ``fuel'';
    (ii) For reciprocating engine powered airplanes, the minimum fuel 
grade;
    (iii) For turbine engine powered airplanes, the permissible fuel 
designations; and
    (iv) For pressure fueling systems, the maximum permissible fueling 
supply pressure and the maximum permissible defueling pressure.
    (2) Oil filler openings must be marked at or near the filler cover 
with the word ``oil''.
    (3) Augmentation fluid filler openings must be marked at or near the 
filler cover to identify the required fluid.
    (c) Emergency exit placards. Each emergency exit placard must meet 
the requirements of Sec. 25.811.
    (d) Doors. Each door that must be used in order to reach any 
required emergency exit must have a suitable placard stating that the 
door is to be latched in the open position during takeoff and landing.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-32, 
37 FR 3972, Feb. 24, 1972; Amdt. 25-38, 41 FR 55468, Dec. 20, 1976; 
Amdt. 25-72, 55 FR 29786, July 20, 1990]



Sec. 25.1561  Safety equipment.

    (a) Each safety equipment control to be operated by the crew in 
emergency, such as controls for automatic liferaft releases, must be 
plainly marked as to its method of operation.

[[Page 538]]

    (b) Each location, such as a locker or compartment, that carries any 
fire extinguishing, signaling, or other life saving equipment must be 
marked accordingly.
    (c) Stowage provisions for required emergency equipment must be 
conspicuously marked to identify the contents and facilitate the easy 
removal of the equipment.
    (d) Each liferaft must have obviously marked operating instructions.
    (e) Approved survival equipment must be marked for identification 
and method of operation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 
43 FR 50598, Oct. 30, 1978]



Sec. 25.1563  Airspeed placard.

    A placard showing the maximum airspeeds for flap extension for the 
takeoff, approach, and landing positions must be installed in clear view 
of each pilot.

                         Airplane Flight Manual



Sec. 25.1581  General.

    (a) Furnishing information. An Airplane Flight Manual must be 
furnished with each airplane, and it must contain the following:
    (1) Information required by Sec. Sec. 25.1583 through 25.1587.
    (2) Other information that is necessary for safe operation because 
of design, operating, or handling characteristics.
    (3) Any limitation, procedure, or other information established as a 
condition of compliance with the applicable noise standards of part 36 
of this chapter.
    (b) Approved information. Each part of the manual listed in 
Sec. Sec. 25.1583 through 25.1587, that is appropriate to the airplane, 
must be furnished, verified, and approved, and must be segregated, 
identified, and clearly distinguished from each unapproved part of that 
manual.
    (c) [Reserved]
    (d) Each Airplane Flight Manual must include a table of contents if 
the complexity of the manual indicates a need for it.

[Amdt. 25-42, 43 FR 2323, Jan. 16, 1978, as amended by Amdt. 25-72, 55 
FR 29786, July 20, 1990]



Sec. 25.1583  Operating limitations.

    (a) Airspeed limitations. The following airspeed limitations and any 
other airspeed limitations necessary for safe operation must be 
furnished:
    (1) The maximum operating limit speed VMO/MMO and a statement that 
this speed limit may not be deliberately exceeded in any regime of 
flight (climb, cruise, or descent) unless a higher speed is authorized 
for flight test or pilot training.
    (2) If an airspeed limitation is based upon compressibility effects, 
a statement to this effect and information as to any symptoms, the 
probable behavior of the airplane, and the recommended recovery 
procedures.
    (3) The maneuvering speed established under Sec. 25.1507 and 
statements, as applicable to the particular design, explaining that:
    (i) Full application of pitch, roll, or yaw controls should be 
confined to speeds below the maneuvering speed; and
    (ii) Rapid and large alternating control inputs, especially in 
combination with large changes in pitch, roll, or yaw, and full control 
inputs in more than one axis at the same time, should be avoided as they 
may result in structural failures at any speed, including below the 
maneuvering speed.
    (4) The flap extended speed VFE and the pertinent flap positions and 
engine powers.
    (5) The landing gear operating speed or speeds, and a statement 
explaining the speeds as defined in Sec. 25.1515(a).
    (6) The landing gear extended speed VLE, if greater than VLO, and a 
statement that this is the maximum speed at which the airplane can be 
safely flown with the landing gear extended.
    (b) Powerplant limitations. The following information must be 
furnished:
    (1) Limitations required by Sec. 25.1521 and Sec. 25.1522.
    (2) Explanation of the limitations, when appropriate.

[[Page 539]]

    (3) Information necessary for marking the instruments required by 
Sec. Sec. 25.1549 through 25.1553.
    (c) Weight and loading distribution. The weight and center of 
gravity limitations established under Sec. 25.1519 must be furnished in 
the Airplane Flight Manual. All of the following information, including 
the weight distribution limitations established under Sec. 25.1519, 
must be presented either in the Airplane Flight Manual or in a separate 
weight and balance control and loading document that is incorporated by 
reference in the Airplane Flight Manual:
    (1) The condition of the airplane and the items included in the 
empty weight as defined in accordance with Sec. 25.29.
    (2) Loading instructions necessary to ensure loading of the airplane 
within the weight and center of gravity limits, and to maintain the 
loading within these limits in flight.
    (3) If certification for more than one center of gravity range is 
requested, the appropriate limitations, with regard to weight and 
loading procedures, for each separate center of gravity range.
    (d) Flight crew. The number and functions of the minimum flight crew 
determined under Sec. 25.1523 must be furnished.
    (e) Kinds of operation. The kinds of operation approved under Sec. 
25.1525 must be furnished.
    (f) Ambient air temperatures and operating altitudes. The extremes 
of the ambient air temperatures and operating altitudes established 
under Sec. 25.1527 must be furnished.
    (g) [Reserved]
    (h) Additional operating limitations. The operating limitations 
established under Sec. 25.1533 must be furnished.
    (i) Maneuvering flight load factors. The positive maneuvering limit 
load factors for which the structure is proven, described in terms of 
accelerations, must be furnished.

[Doc. No. 5066, 29 FR 1891, Dec. 24, 1964, as amended by Amdt. 25-38, 41 
FR 55468, Dec, 20, 1976; Amdt. 25-42, 43 FR 2323, Jan. 16, 1978; Amdt. 
25-46, 43 FR 50598, Oct. 30, 1978; Amdt. 25-72, 55 FR 29787, July 20, 
1990; Amdt. 25-105, 66 FR 34024, June 26, 2001; 75 FR 49818, Aug. 16, 
2010]



Sec. 25.1585  Operating procedures.

    (a) Operating procedures must be furnished for--
    (1) Normal procedures peculiar to the particular type or model 
encountered in connection with routine operations;
    (2) Non-normal procedures for malfunction cases and failure 
conditions involving the use of special systems or the alternative use 
of regular systems; and
    (3) Emergency procedures for foreseeable but unusual situations in 
which immediate and precise action by the crew may be expected to 
substantially reduce the risk of catastrophe.
    (b) Information or procedures not directly related to airworthiness 
or not under the control of the crew, must not be included, nor must any 
procedure that is accepted as basic airmanship.
    (c) Information identifying each operating condition in which the 
fuel system independence prescribed in Sec. 25.953 is necessary for 
safety must be furnished, together with instructions for placing the 
fuel system in a configuration used to show compliance with that 
section.
    (d) The buffet onset envelopes, determined under Sec. 25.251 must 
be furnished. The buffet onset envelopes presented may reflect the 
center of gravity at which the airplane is normally loaded during cruise 
if corrections for the effect of different center of gravity locations 
are furnished.
    (e) Information must be furnished that indicates that when the fuel 
quantity indicator reads ``zero'' in level flight, any fuel remaining in 
the fuel tank cannot be used safely in flight.
    (f) Information on the total quantity of usable fuel for each fuel 
tank must be furnished.

[Doc. No. 2000-8511, 66 FR 34024, June 26, 2001]



Sec. 25.1587  Performance information.

    (a) Each Airplane Flight Manual must contain information to permit 
conversion of the indicated temperature to free air temperature if other 
than a free air temperature indicator is used to comply with the 
requirements of Sec. 25.1303(a)(1).
    (b) Each Airplane Flight Manual must contain the performance 
information computed under the applicable

[[Page 540]]

provisions of this part (including Sec. Sec. 25.115, 25.123, and 25.125 
for the weights, altitudes, temperatures, wind components, and runway 
gradients, as applicable) within the operational limits of the airplane, 
and must contain the following:
    (1) In each case, the conditions of power, configuration, and 
speeds, and the procedures for handling the airplane and any system 
having a significant effect on the performance information.
    (2) VSR determined in accordance with Sec. 25.103.
    (3) The following performance information (determined by 
extrapolation and computed for the range of weights between the maximum 
landing weight and the maximum takeoff weight):
    (i) Climb in the landing configuration.
    (ii) Climb in the approach configuration.
    (iii) Landing distance.
    (4) Procedures established under Sec. 25.101(f) and (g) that are 
related to the limitations and information required by Sec. 25.1533 and 
by this paragraph (b) in the form of guidance material, including any 
relevant limitations or information.
    (5) An explanation of significant or unusual flight or ground 
handling characteristics of the airplane.
    (6) Corrections to indicated values of airspeed, altitude, and 
outside air temperature.
    (7) An explanation of operational landing runway length factors 
included in the presentation of the landing distance, if appropriate.

[Doc. No. 2000-8511, 66 FR 34024, June 26, 2001, as amended by Amdt. 25-
108, 67 FR 70828, Nov. 26, 2002]



       Subpart H_Electrical Wiring Interconnection Systems (EWIS)

    Source: Docket No. FAA-2004-18379, 72 FR 63406, Nov. 8, 2007, unless 
otherwise noted.



Sec. 25.1701  Definition.

    (a) As used in this chapter, electrical wiring interconnection 
system (EWIS) means any wire, wiring device, or combination of these, 
including termination devices, installed in any area of the airplane for 
the purpose of transmitting electrical energy, including data and 
signals, between two or more intended termination points. This includes:
    (1) Wires and cables.
    (2) Bus bars.
    (3) The termination point on electrical devices, including those on 
relays, interrupters, switches, contactors, terminal blocks and circuit 
breakers, and other circuit protection devices.
    (4) Connectors, including feed-through connectors.
    (5) Connector accessories.
    (6) Electrical grounding and bonding devices and their associated 
connections.
    (7) Electrical splices.
    (8) Materials used to provide additional protection for wires, 
including wire insulation, wire sleeving, and conduits that have 
electrical termination for the purpose of bonding.
    (9) Shields or braids.
    (10) Clamps and other devices used to route and support the wire 
bundle.
    (11) Cable tie devices.
    (12) Labels or other means of identification.
    (13) Pressure seals.
    (14) EWIS components inside shelves, panels, racks, junction boxes, 
distribution panels, and back-planes of equipment racks, including, but 
not limited to, circuit board back-planes, wire integration units, and 
external wiring of equipment.
    (b) Except for the equipment indicated in paragraph (a)(14) of this 
section, EWIS components inside the following equipment, and the 
external connectors that are part of that equipment, are excluded from 
the definition in paragraph (a) of this section:
    (1) Electrical equipment or avionics that are qualified to 
environmental conditions and testing procedures when those conditions 
and procedures are--
    (i) Appropriate for the intended function and operating environment, 
and
    (ii) Acceptable to the FAA.
    (2) Portable electrical devices that are not part of the type design 
of the airplane. This includes personal entertainment devices and laptop 
computers.
    (3) Fiber optics.

[[Page 541]]



Sec. 25.1703  Function and installation: EWIS.

    (a) Each EWIS component installed in any area of the aircraft must:
    (1) Be of a kind and design appropriate to its intended function.
    (2) Be installed according to limitations specified for the EWIS 
components.
    (3) Perform the function for which it was intended without degrading 
the airworthiness of the airplane.
    (4) Be designed and installed in a way that will minimize mechanical 
strain.
    (b) Selection of wires must take into account known characteristics 
of the wire in relation to each installation and application to minimize 
the risk of wire damage, including any arc tracking phenomena.
    (c) The design and installation of the main power cables (including 
generator cables) in the fuselage must allow for a reasonable degree of 
deformation and stretching without failure.
    (d) EWIS components located in areas of known moisture accumulation 
must be protected to minimize any hazardous effects due to moisture.



Sec. 25.1705  Systems and functions: EWIS.

    (a) EWIS associated with any system required for type certification 
or by operating rules must be considered an integral part of that system 
and must be considered in showing compliance with the applicable 
requirements for that system.
    (b) For systems to which the following rules apply, the components 
of EWIS associated with those systems must be considered an integral 
part of that system or systems and must be considered in showing 
compliance with the applicable requirements for that system.
    (1) Sec. 25.773(b)(2) Pilot compartment view.
    (2) Sec. 25.981 Fuel tank ignition prevention.
    (3) Sec. 25.1165 Engine ignition systems.
    (4) Sec. 25.1310 Power source capacity and distribution.
    (5) Sec. 25.1316 System lightning protection.
    (6) Sec. 25.1331(a)(2) Instruments using a power supply.
    (7) Sec. 25.1351 General.
    (8) Sec. 25.1355 Distribution system.
    (9) Sec. 25.1360 Precautions against injury.
    (10) Sec. 25.1362 Electrical supplies for emergency conditions.
    (11) Sec. 25.1365 Electrical appliances, motors, and transformers.
    (12) Sec. 25.1431(c) and (d) Electronic equipment.



Sec. 25.1707  System separation: EWIS.

    (a) Each EWIS must be designed and installed with adequate physical 
separation from other EWIS and airplane systems so that an EWIS 
component failure will not create a hazardous condition. Unless 
otherwise stated, for the purposes of this section, adequate physical 
separation must be achieved by separation distance or by a barrier that 
provides protection equivalent to that separation distance.
    (b) Each EWIS must be designed and installed so that any electrical 
interference likely to be present in the airplane will not result in 
hazardous effects upon the airplane or its systems.
    (c) Wires and cables carrying heavy current, and their associated 
EWIS components, must be designed and installed to ensure adequate 
physical separation and electrical isolation so that damage to circuits 
associated with essential functions will be minimized under fault 
conditions.
    (d) Each EWIS associated with independent airplane power sources or 
power sources connected in combination must be designed and installed to 
ensure adequate physical separation and electrical isolation so that a 
fault in any one airplane power source EWIS will not adversely affect 
any other independent power sources. In addition:
    (1) Airplane independent electrical power sources must not share a 
common ground terminating location.
    (2) Airplane system static grounds must not share a common ground 
terminating location with any of the airplane's independent electrical 
power sources.
    (e) Except to the extent necessary to provide electrical connection 
to the fuel systems components, the EWIS must be designed and installed 
with adequate physical separation from fuel lines and other fuel system 
components, so that:

[[Page 542]]

    (1) An EWIS component failure will not create a hazardous condition.
    (2) Any fuel leakage onto EWIS components will not create a 
hazardous condition.
    (f) Except to the extent necessary to provide electrical connection 
to the hydraulic systems components, EWIS must be designed and installed 
with adequate physical separation from hydraulic lines and other 
hydraulic system components, so that:
    (1) An EWIS component failure will not create a hazardous condition.
    (2) Any hydraulic fluid leakage onto EWIS components will not create 
a hazardous condition.
    (g) Except to the extent necessary to provide electrical connection 
to the oxygen systems components, EWIS must be designed and installed 
with adequate physical separation from oxygen lines and other oxygen 
system components, so that an EWIS component failure will not create a 
hazardous condition.
    (h) Except to the extent necessary to provide electrical connection 
to the water/waste systems components, EWIS must be designed and 
installed with adequate physical separation from water/waste lines and 
other water/waste system components, so that:
    (1) An EWIS component failure will not create a hazardous condition.
    (2) Any water/waste leakage onto EWIS components will not create a 
hazardous condition.
    (i) EWIS must be designed and installed with adequate physical 
separation between the EWIS and flight or other mechanical control 
systems cables and associated system components, so that:
    (1) Chafing, jamming, or other interference are prevented.
    (2) An EWIS component failure will not create a hazardous condition.
    (3) Failure of any flight or other mechanical control systems cables 
or systems components will not damage the EWIS and create a hazardous 
condition.
    (j) EWIS must be designed and installed with adequate physical 
separation between the EWIS components and heated equipment, hot air 
ducts, and lines, so that:
    (1) An EWIS component failure will not create a hazardous condition.
    (2) Any hot air leakage or heat generated onto EWIS components will 
not create a hazardous condition.
    (k) For systems for which redundancy is required, by certification 
rules, by operating rules, or as a result of the assessment required by 
Sec. 25.1709, EWIS components associated with those systems must be 
designed and installed with adequate physical separation.
    (l) Each EWIS must be designed and installed so there is adequate 
physical separation between it and other aircraft components and 
aircraft structure, and so that the EWIS is protected from sharp edges 
and corners, to minimize potential for abrasion/chafing, vibration 
damage, and other types of mechanical damage.



Sec. 25.1709  System safety: EWIS.

    Each EWIS must be designed and installed so that:
    (a) Each catastrophic failure condition--
    (1) Is extremely improbable; and
    (2) Does not result from a single failure.
    (b) Each hazardous failure condition is extremely remote.



Sec. 25.1711  Component identification: EWIS.

    (a) EWIS components must be labeled or otherwise identified using a 
consistent method that facilitates identification of the EWIS component, 
its function, and its design limitations, if any.
    (b) For systems for which redundancy is required, by certification 
rules, by operating rules, or as a result of the assessment required by 
Sec. 25.1709, EWIS components associated with those systems must be 
specifically identified with component part number, function, and 
separation requirement for bundles.
    (1) The identification must be placed along the wire, cable, or wire 
bundle at appropriate intervals and in areas of the airplane where it is 
readily visible to maintenance, repair, or alteration personnel.

[[Page 543]]

    (2) If an EWIS component cannot be marked physically, then other 
means of identification must be provided.
    (c) The identifying markings required by paragraphs (a) and (b) of 
this section must remain legible throughout the expected service life of 
the EWIS component.
    (d) The means used for identifying each EWIS component as required 
by this section must not have an adverse effect on the performance of 
that component throughout its expected service life.
    (e) Identification for EWIS modifications to the type design must be 
consistent with the identification scheme of the original type design.



Sec. 25.1713  Fire protection: EWIS.

    (a) All EWIS components must meet the applicable fire and smoke 
protection requirements of Sec. 25.831(c) of this part.
    (b) EWIS components that are located in designated fire zones and 
are used during emergency procedures must be fire resistant.
    (c) Insulation on electrical wire and electrical cable, and 
materials used to provide additional protection for the wire and cable, 
installed in any area of the airplane, must be self-extinguishing when 
tested in accordance with the applicable portions of Appendix F, part I, 
of 14 CFR part 25.



Sec. 25.1715  Electrical bonding and protection against static
electricity: EWIS.

    (a) EWIS components used for electrical bonding and protection 
against static electricity must meet the requirements of Sec. 25.899.
    (b) On airplanes having grounded electrical systems, electrical 
bonding provided by EWIS components must provide an electrical return 
path capable of carrying both normal and fault currents without creating 
a shock hazard or damage to the EWIS components, other airplane system 
components, or airplane structure.



Sec. 25.1717  Circuit protective devices: EWIS.

    Electrical wires and cables must be designed and installed so they 
are compatible with the circuit protection devices required by Sec. 
25.1357, so that a fire or smoke hazard cannot be created under 
temporary or continuous fault conditions.



Sec. 25.1719  Accessibility provisions: EWIS.

    Access must be provided to allow inspection and replacement of any 
EWIS component as necessary for continued airworthiness.



Sec. 25.1721  Protection of EWIS.

    (a) No cargo or baggage compartment may contain any EWIS whose 
damage or failure may affect safe operation, unless the EWIS is 
protected so that:
    (1) It cannot be damaged by movement of cargo or baggage in the 
compartment.
    (2) Its breakage or failure will not create a fire hazard.
    (b) EWIS must be designed and installed to minimize damage and risk 
of damage to EWIS by movement of people in the airplane during all 
phases of flight, maintenance, and servicing.
    (c) EWIS must be designed and installed to minimize damage and risk 
of damage to EWIS by items carried onto the aircraft by passengers or 
cabin crew.



Sec. 25.1723  Flammable fluid fire protection: EWIS.

    EWIS components located in each area where flammable fluid or vapors 
might escape by leakage of a fluid system must be considered a potential 
ignition source and must meet the requirements of Sec. 25.863.



Sec. 25.1725  Powerplants: EWIS.

    (a) EWIS associated with any powerplant must be designed and 
installed so that the failure of an EWIS component will not prevent the 
continued safe operation of the remaining powerplants or require 
immediate action by any crewmember for continued safe operation, in 
accordance with the requirements of Sec. 25.903(b).
    (b) Design precautions must be taken to minimize hazards to the 
airplane due to EWIS damage in the event of a powerplant rotor failure 
or a fire originating within the powerplant that burns through the 
powerplant case, in

[[Page 544]]

accordance with the requirements of Sec. 25.903(d)(1).



Sec. 25.1727  Flammable fluid shutoff means: EWIS.

    EWIS associated with each flammable fluid shutoff means and control 
must be fireproof or must be located and protected so that any fire in a 
fire zone will not affect operation of the flammable fluid shutoff 
means, in accordance with the requirements of Sec. 25.1189.



Sec. 25.1729  Instructions for Continued Airworthiness: EWIS.

    The applicant must prepare Instructions for Continued Airworthiness 
applicable to EWIS in accordance with Appendix H sections H25.4 and 
H25.5 to this part that are approved by the FAA.



Sec. 25.1731  Powerplant and APU fire detector system: EWIS.

    (a) EWIS that are part of each fire or overheat detector system in a 
fire zone must be fire-resistant.
    (b) No EWIS component of any fire or overheat detector system for 
any fire zone may pass through another fire zone, unless:
    (1) It is protected against the possibility of false warnings 
resulting from fires in zones through which it passes; or
    (2) Each zone involved is simultaneously protected by the same 
detector and extinguishing system.
    (c) EWIS that are part of each fire or overheat detector system in a 
fire zone must meet the requirements of Sec. 25.1203.



Sec. 25.1733  Fire detector systems, general: EWIS.

    EWIS associated with any installed fire protection system, including 
those required by Sec. Sec. 25.854 and 25.858, must be considered an 
integral part of the system in showing compliance with the applicable 
requirements for that system.



             Subpart I_Special Federal Aviation Regulations

    Source: Docket No. FAA-2011-0186, Amdt. 25-133, 76 FR 12555, Mar. 8, 
2011, unless otherwise noted.



Sec. 25.1801  SFAR No. 111--Lavatory Oxygen Systems.

    The requirements of Sec. 121.1500 of this chapter also apply to 
this part.

[[Page 545]]



                       Sec. Appendix A to Part 25
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                       Sec. Appendix B to Part 25
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                       Sec. Appendix C to Part 25

                  Part I--Atmospheric Icing Conditions

    (a) Continuous maximum icing. The maximum continuous intensity of 
atmospheric icing conditions (continuous maximum icing) is defined by 
the variables of the cloud liquid water content, the mean effective 
diameter of the cloud droplets, the ambient air temperature, and the 
interrelationship of these three variables as shown in figure 1 of this 
appendix. The limiting icing envelope in terms of altitude and 
temperature is given in figure 2 of this appendix. The inter-
relationship of cloud liquid water content with drop diameter and 
altitude is determined from figures 1 and 2. The cloud liquid water 
content for continuous maximum icing conditions of a horizontal extent, 
other than 17.4 nautical miles, is determined by the value of liquid 
water content of figure 1, multiplied by the appropriate factor from 
figure 3 of this appendix.
    (b) Intermittent maximum icing. The intermittent maximum intensity 
of atmospheric icing conditions (intermittent maximum icing) is defined 
by the variables of the cloud liquid water content, the mean effective 
diameter of the cloud droplets, the ambient air temperature, and the 
interrelationship of these three variables as shown in figure 4 of this 
appendix. The limiting icing envelope in terms of altitude and 
temperature is given in figure 5 of this appendix. The inter-
relationship of cloud liquid water content with drop diameter and 
altitude is determined from figures 4 and 5. The cloud liquid water 
content for intermittent maximum icing conditions of a horizontal 
extent, other than 2.6 nautical miles, is determined by the value of 
cloud liquid water content of figure 4 multiplied by the appropriate 
factor in figure 6 of this appendix.

[[Page 553]]

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    (c) Takeoff maximum icing. The maximum intensity of atmospheric 
icing conditions for takeoff (takeoff maximum icing) is defined by the 
cloud liquid water content of 0.35 g/m3, the mean effective diameter of 
the cloud droplets of 20 microns, and the ambient air temperature at 
ground level of minus 9 degrees Celsius (-9 [deg]C). The takeoff maximum 
icing conditions extend from ground level to a height of 1,500 feet 
above the level of the takeoff surface.

[[Page 559]]

 Part II--Airframe Ice Accretions for Showing Compliance With Subpart B.

    (a) Ice accretions--General. The most critical ice accretion in 
terms of airplane performance and handling qualities for each flight 
phase must be used to show compliance with the applicable airplane 
performance and handling requirements in icing conditions of subpart B 
of this part. Applicants must demonstrate that the full range of 
atmospheric icing conditions specified in part I of this appendix have 
been considered, including the mean effective drop diameter, liquid 
water content, and temperature appropriate to the flight conditions (for 
example, configuration, speed, angle-of-attack, and altitude). The ice 
accretions for each flight phase are defined as follows:
    (1) Takeoff ice is the most critical ice accretion on unprotected 
surfaces and any ice accretion on the protected surfaces appropriate to 
normal ice protection system operation, occurring between liftoff and 
400 feet above the takeoff surface, assuming accretion starts at liftoff 
in the takeoff maximum icing conditions of part I, paragraph (c) of this 
appendix.
    (2) Final takeoff ice is the most critical ice accretion on 
unprotected surfaces, and any ice accretion on the protected surfaces 
appropriate to normal ice protection system operation, between 400 feet 
and either 1,500 feet above the takeoff surface, or the height at which 
the transition from the takeoff to the en route configuration is 
completed and VFTO is reached, whichever is higher. Ice 
accretion is assumed to start at liftoff in the takeoff maximum icing 
conditions of part I, paragraph (c) of this appendix.
    (3) En route ice is the critical ice accretion on the unprotected 
surfaces, and any ice accretion on the protected surfaces appropriate to 
normal ice protection system operation, during the en route phase.
    (4) Holding ice is the critical ice accretion on the unprotected 
surfaces, and any ice accretion on the protected surfaces appropriate to 
normal ice protection system operation, during the holding flight phase.
    (5) Approach ice is the critical ice accretion on the unprotected 
surfaces, and any ice accretion on the protected surfaces appropriate to 
normal ice protection system operation following exit from the holding 
flight phase and transition to the most critical approach configuration.
    (6) Landing ice is the critical ice accretion on the unprotected 
surfaces, and any ice accretion on the protected surfaces appropriate to 
normal ice protection system operation following exit from the approach 
flight phase and transition to the final landing configuration.
    (b) In order to reduce the number of ice accretions to be considered 
when demonstrating compliance with the requirements of Sec. 25.21(g), 
any of the ice accretions defined in paragraph (a) of this section may 
be used for any other flight phase if it is shown to be more critical 
than the specific ice accretion defined for that flight phase. 
Configuration differences and their effects on ice accretions must be 
taken into account.
    (c) The ice accretion that has the most adverse effect on handling 
qualities may be used for airplane performance tests provided any 
difference in performance is conservatively taken into account.
    (d) For both unprotected and protected parts, the ice accretion for 
the takeoff phase may be determined by calculation, assuming the takeoff 
maximum icing conditions defined in appendix C, and assuming that:
    (1) Airfoils, control surfaces and, if applicable, propellers are 
free from frost, snow, or ice at the start of the takeoff;
    (2) The ice accretion starts at liftoff;
    (3) The critical ratio of thrust/power-to-weight;
    (4) Failure of the critical engine occurs at VEF; and
    (5) Crew activation of the ice protection system is in accordance 
with a normal operating procedure provided in the Airplane Flight 
Manual, except that after beginning the takeoff roll, it must be assumed 
that the crew takes no action to activate the ice protection system 
until the airplane is at least 400 feet above the takeoff surface.
    (e) The ice accretion before the ice protection system has been 
activated and is performing its intended function is the critical ice 
accretion formed on the unprotected and normally protected surfaces 
before activation and effective operation of the ice protection system 
in continuous maximum atmospheric icing conditions. This ice accretion 
only applies in showing compliance to Sec. Sec. 25.143(j) and 
25.207(h), and 25.207(i).

[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 25-121, 
72 FR 44669, Aug. 8, 2007; 72 FR 50467, Aug. 31, 2007; Amdt. 25-129, 74 
FR 38340, Aug. 3, 2009]

    Effective Date Note: At 79 FR 65528, Nov. 4, 2014, Part 25, Appendix 
C, Part II was amended by revising paragraph (a)(1), the second sentence 
of paragraph (a)(2), and paragraph (d)(2), effective Jan. 5, 2015. For 
the convenience of the user, the revised text is set forth as follows:



                       Sec. Appendix C to Part 25

                                * * * * *

 PART II--AIRFRAME ICE ACCRETIONS FOR SHOWING COMPLIANCE WITH SUBPART B

    (a) * * *

[[Page 560]]

    (1) Takeoff ice is the most critical ice accretion on unprotected 
surfaces and any ice accretion on the protected surfaces appropriate to 
normal ice protection system operation, occurring between the end of the 
takeoff distance and 400 feet above the takeoff surface, assuming 
accretion starts at the end of the takeoff distance in the takeoff 
maximum icing conditions defined in part I of this Appendix.
    (2) * * * Ice accretion is assumed to start at the end of the 
takeoff distance in the takeoff maximum icing conditions of part I, 
paragraph (c) of this Appendix.

                                * * * * *

    (d) * * *
    (2) The ice accretion starts at the end of the takeoff distance.

                                * * * * *



                       Sec. Appendix D to Part 25

    Criteria for determining minimum flight crew. The following are 
considered by the Agency in determining the minimum flight crew under 
Sec. 25.1523:
    (a) Basic workload functions. The following basic workload functions 
are considered:
    (1) Flight path control.
    (2) Collision avoidance.
    (3) Navigation.
    (4) Communications.
    (5) Operation and monitoring of aircraft engines and systems.
    (6) Command decisions.
    (b) Workload factors. The following workload factors are considered 
significant when analyzing and demonstrating workload for minimum flight 
crew determination:
    (1) The accessibility, ease, and simplicity of operation of all 
necessary flight, power, and equipment controls, including emergency 
fuel shutoff valves, electrical controls, electronic controls, 
pressurization system controls, and engine controls.
    (2) The accessibility and conspicuity of all necessary instruments 
and failure warning devices such as fire warning, electrical system 
malfunction, and other failure or caution indicators. The extent to 
which such instruments or devices direct the proper corrective action is 
also considered.
    (3) The number, urgency, and complexity of operating procedures with 
particular consideration given to the specific fuel management schedule 
imposed by center of gravity, structural or other considerations of an 
airworthiness nature, and to the ability of each engine to operate at 
all times from a single tank or source which is automatically 
replenished if fuel is also stored in other tanks.
    (4) The degree and duration of concentrated mental and physical 
effort involved in normal operation and in diagnosing and coping with 
malfunctions and emergencies.
    (5) The extent of required monitoring of the fuel, hydraulic, 
pressurization, electrical, electronic, deicing, and other systems while 
en route.
    (6) The actions requiring a crewmember to be unavailable at his 
assigned duty station, including: observation of systems, emergency 
operation of any control, and emergencies in any compartment.
    (7) The degree of automation provided in the aircraft systems to 
afford (after failures or malfunctions) automatic crossover or isolation 
of difficulties to minimize the need for flight crew action to guard 
against loss of hydraulic or electric power to flight controls or to 
other essential systems.
    (8) The communications and navigation workload.
    (9) The possibility of increased workload associated with any 
emergency that may lead to other emergencies.
    (10) Incapacitation of a flight crewmember whenever the applicable 
operating rule requires a minimum flight crew of at least two pilots.
    (c) Kind of operation authorized. The determination of the kind of 
operation authorized requires consideration of the operating rules under 
which the airplane will be operated. Unless an applicant desires 
approval for a more limited kind of operation. It is assumed that each 
airplane certificated under this Part will operate under IFR conditions.

[Amdt. 25-3, 30 FR 6067, Apr. 29, 1965]



                       Sec. Appendix E to Part 25

   I--Limited Weight Credit For Airplanes Equipped With Standby Power

    (a) Each applicant for an increase in the maximum certificated 
takeoff and landing weights of an airplane equipped with a type-
certificated standby power rocket engine may obtain an increase as 
specified in paragraph (b) if--
    (1) The installation of the rocket engine has been approved and it 
has been established by flight test that the rocket engine and its 
controls can be operated safely and reliably at the increase in maximum 
weight; and
    (2) The Airplane Flight Manual, or the placard, markings or manuals 
required in place thereof, set forth in addition to any other operating 
limitations the Administrator may require, the increased weight approved 
under this regulation and a prohibition against the operation of the 
airplane at the approved increased weight when--
    (i) The installed standby power rocket engines have been stored or 
installed in excess of the time limit established by the manufacturer of 
the rocket engine (usually stenciled on the engine casing); or

[[Page 561]]

    (ii) The rocket engine fuel has been expended or discharged.
    (b) The currently approved maximum takeoff and landing weights at 
which an airplane is certificated without a standby power rocket engine 
installation may be increased by an amount that does not exceed any of 
the following:
    (1) An amount equal in pounds to 0.014 IN, where I is the maximum 
usable impulse in pounds-seconds available from each standby power 
rocket engine and N is the number of rocket engines installed.
    (2) An amount equal to 5 percent of the maximum certificated weight 
approved in accordance with the applicable airworthiness regulations 
without standby power rocket engines installed.
    (3) An amount equal to the weight of the rocket engine installation.
    (4) An amount that, together with the currently approved maximum 
weight, would equal the maximum structural weight established for the 
airplane without standby rocket engines installed.

 II--Performance Credit for Transport Category Airplanes Equipped With 
                              Standby Power

    The Administrator may grant performance credit for the use of 
standby power on transport category airplanes. However, the performance 
credit applies only to the maximum certificated takeoff and landing 
weights, the takeoff distance, and the takeoff paths, and may not exceed 
that found by the Administrator to result in an overall level of safety 
in the takeoff, approach, and landing regimes of flight equivalent to 
that prescribed in the regulations under which the airplane was 
originally certificated without standby power. For the purposes of this 
appendix, ``standby power'' is power or thrust, or both, obtained from 
rocket engines for a relatively short period and actuated only in cases 
of emergency. The following provisions apply:
    (1) Takeoff; general. The takeoff data prescribed in paragraphs (2) 
and (3) of this appendix must be determined at all weights and 
altitudes, and at ambient temperatures if applicable, at which 
performance credit is to be applied.
    (2) Takeoff path.
    (a) The one-engine-inoperative takeoff path with standby power in 
use must be determined in accordance with the performance requirements 
of the applicable airworthiness regulations.
    (b) The one-engine-inoperative takeoff path (excluding that part 
where the airplane is on or just above the takeoff surface) determined 
in accordance with paragraph (a) of this section must lie above the one-
engine-inoperative takeoff path without standby power at the maximum 
takeoff weight at which all of the applicable air-worthiness 
requirements are met. For the purpose of this comparison, the flight 
path is considered to extend to at least a height of 400 feet above the 
takeoff surface.
    (c) The takeoff path with all engines operating, but without the use 
of standby power, must reflect a conservatively greater overall level of 
performance than the one-engine-inoperative takeoff path established in 
accordance with paragraph (a) of this section. The margin must be 
established by the Administrator to insure safe day-to-day operations, 
but in no case may it be less than 15 percent. The all-engines-operating 
takeoff path must be determined by a procedure consistent with that 
established in complying with paragraph (a) of this section.
    (d) For reciprocating-engine-powered airplanes, the takeoff path to 
be scheduled in the Airplane Flight Manual must represent the one-
engine-operative takeoff path determined in accordance with paragraph 
(a) of this section and modified to reflect the procedure (see paragraph 
(6)) established by the applicant for flap retraction and attainment of 
the en route speed. The scheduled takeoff path must have a positive 
slope at all points of the airborne portion and at no point must it lie 
above the takeoff path specified in paragraph (a) of this section.
    (3) Takeoff distance. The takeoff distance must be the horizontal 
distance along the one-engine-inoperative take off path determined in 
accordance with paragraph (2)(a) from the start of the takeoff to the 
point where the airplane attains a height of 50 feet above the takeoff 
surface for reciprocating-engine-powered airplanes and a height of 35 
feet above the takeoff surface for turbine-powered airplanes.
    (4) Maximum certificated takeoff weights. The maximum certificated 
takeoff weights must be determined at all altitudes, and at ambient 
temperatures, if applicable, at which performance credit is to be 
applied and may not exceed the weights established in compliance with 
paragraphs (a) and (b) of this section.
    (a) The conditions of paragraphs (2)(b) through (d) must be met at 
the maximum certificated takeoff weight.
    (b) Without the use of standby power, the airplane must meet all of 
the en route requirements of the applicable airworthiness regulations 
under which the airplane was originally certificated. In addition, 
turbine-powered airplanes without the use of standby power must meet the 
final takeoff climb requirements prescribed in the applicable 
airworthiness regulations.
    (5) Maximum certificated landing weights.
    (a) The maximum certificated landing weights (one-engine-inoperative 
approach and all-engine-operating landing climb) must be determined at 
all altitudes, and at ambient temperatures if applicable, at which 
performance credit is to be applied and must not exceed that established 
in compliance with paragraph (b) of this section.

[[Page 562]]

    (b) The flight path, with the engines operating at the power or 
thrust, or both, appropriate to the airplane configuration and with 
standby power in use, must lie above the flight path without standby 
power in use at the maximum weight at which all of the applicable 
airworthiness requirements are met. In addition, the flight paths must 
comply with subparagraphs (i) and (ii) of this paragraph.
    (i) The flight paths must be established without changing the 
appropriate airplane configuration.
    (ii) The flight paths must be carried out for a minimum height of 
400 feet above the point where standby power is actuated.
    (6) Airplane configuration, speed, and power and thrust; general. 
Any change in the airplane's configuration, speed, and power or thrust, 
or both, must be made in accordance with the procedures established by 
the applicant for the operation of the airplane in service and must 
comply with paragraphs (a) through (c) of this section. In addition, 
procedures must be established for the execution of balked landings and 
missed approaches.
    (a) The Administrator must find that the procedure can be 
consistently executed in service by crews of average skill.
    (b) The procedure may not involve methods or the use of devices 
which have not been proven to be safe and reliable.
    (c) Allowances must be made for such time delays in the execution of 
the procedures as may be reasonably expected to occur during service.
    (7) Installation and operation; standby power. The standby power 
unit and its installation must comply with paragraphs (a) and (b) of 
this section.
    (a) The standby power unit and its installation must not adversely 
affect the safety of the airplane.
    (b) The operation of the standby power unit and its control must 
have proven to be safe and reliable.

[Amdt. 25-6, 30 FR 8468, July 2, 1965]



                       Sec. Appendix F to Part 25

 Part I--Test Criteria and Procedures for Showing Compliance with Sec. 
                        25.853, or Sec. 25.855.

    (a) Material test criteria--(1) Interior compartments occupied by 
crew or passengers. (i) Interior ceiling panels, interior wall panels, 
partitions, galley structure, large cabinet walls, structural flooring, 
and materials used in the construction of stowage compartments (other 
than underseat stowage compartments and compartments for stowing small 
items such as magazines and maps) must be self-extinguishing when tested 
vertically in accordance with the applicable portions of part I of this 
appendix. The average burn length may not exceed 6 inches and the 
average flame time after removal of the flame source may not exceed 15 
seconds. Drippings from the test specimen may not continue to flame for 
more than an average of 3 seconds after falling.
    (ii) Floor covering, textiles (including draperies and upholstery), 
seat cushions, padding, decorative and nondecorative coated fabrics, 
leather, trays and galley furnishings, electrical conduit, air ducting, 
joint and edge covering, liners of Class B and E cargo or baggage 
compartments, floor panels of Class B, C, D, or E cargo or baggage 
compartments, cargo covers and transparencies, molded and thermoformed 
parts, air ducting joints, and trim strips (decorative and chafing), 
that are constructed of materials not covered in subparagraph (iv) 
below, must be self-extinguishing when tested vertically in accordance 
with the applicable portions of part I of this appendix or other 
approved equivalent means. The average burn length may not exceed 8 
inches, and the average flame time after removal of the flame source may 
not exceed 15 seconds. Drippings from the test specimen may not continue 
to flame for more than an average of 5 seconds after falling.
    (iii) Motion picture film must be safety film meeting the Standard 
Specifications for Safety Photographic Film PHI.25 (available from the 
American National Standards Institute, 1430 Broadway, New York, NY 
10018). If the film travels through ducts, the ducts must meet the 
requirements of subparagraph (ii) of this paragraph.
    (iv) Clear plastic windows and signs, parts constructed in whole or 
in part of elastomeric materials, edge lighted instrument assemblies 
consisting of two or more instruments in a common housing, seat belts, 
shoulder harnesses, and cargo and baggage tiedown equipment, including 
containers, bins, pallets, etc., used in passenger or crew compartments, 
may not have an average burn rate greater than 2.5 inches per minute 
when tested horizontally in accordance with the applicable portions of 
this appendix.
    (v) Except for small parts (such as knobs, handles, rollers, 
fasteners, clips, grommets, rub strips, pulleys, and small electrical 
parts) that would not contribute significantly to the propagation of a 
fire and for electrical wire and cable insulation, materials in items 
not specified in paragraphs (a)(1)(i), (ii), (iii), or (iv) of part I of 
this appendix may not have a burn rate greater than 4.0 inches per 
minute when tested horizontally in accordance with the applicable 
portions of this appendix.
    (2) Cargo and baggage compartments not occupied by crew or 
passengers.
    (i) [Reserved]
    (ii) A cargo or baggage compartment defined in Sec. 25.857 as Class 
B or E must have a liner constructed of materials that meet the 
requirements of paragraph (a)(1)(ii) of part I

[[Page 563]]

of this appendix and separated from the airplane structure (except for 
attachments). In addition, such liners must be subjected to the 45 
degree angle test. The flame may not penetrate (pass through) the 
material during application of the flame or subsequent to its removal. 
The average flame time after removal of the flame source may not exceed 
15 seconds, and the average glow time may not exceed 10 seconds.
    (iii) A cargo or baggage compartment defined in Sec. 25.857 as 
Class B, C, D, or E must have floor panels constructed of materials 
which meet the requirements of paragraph (a)(1)(ii) of part I of this 
appendix and which are separated from the airplane structure (except for 
attachments). Such panels must be subjected to the 45 degree angle test. 
The flame may not penetrate (pass through) the material during 
application of the flame or subsequent to its removal. The average flame 
time after removal of the flame source may not exceed 15 seconds, and 
the average glow time may not exceed 10 seconds.
    (iv) Insulation blankets and covers used to protect cargo must be 
constructed of materials that meet the requirements of paragraph 
(a)(1)(ii) of part I of this appendix. Tiedown equipment (including 
containers, bins, and pallets) used in each cargo and baggage 
compartment must be constructed of materials that meet the requirements 
of paragraph (a)(1)(v) of part I of this appendix.
    (3) Electrical system components. Insulation on electrical wire or 
cable installed in any area of the fuselage must be self-extinguishing 
when subjected to the 60 degree test specified in part I of this 
appendix. The average burn length may not exceed 3 inches, and the 
average flame time after removal of the flame source may not exceed 30 
seconds. Drippings from the test specimen may not continue to flame for 
more than an average of 3 seconds after falling.
    (b) Test Procedures--(1) Conditioning. Specimens must be conditioned 
to 70 5 F., and at 50 percent 5 percent relative humidity until moisture equilibrium 
is reached or for 24 hours. Each specimen must remain in the 
conditioning environment until it is subjected to the flame.
    (2) Specimen configuration. Except for small parts and electrical 
wire and cable insulation, materials must be tested either as section 
cut from a fabricated part as installed in the airplane or as a specimen 
simulating a cut section, such as a specimen cut from a flat sheet of 
the material or a model of the fabricated part. The specimen may be cut 
from any location in a fabricated part; however, fabricated units, such 
as sandwich panels, may not be separated for test. Except as noted 
below, the specimen thickness must be no thicker than the minimum 
thickness to be qualified for use in the airplane. Test specimens of 
thick foam parts, such as seat cushions, must be \1/2\-inch in 
thickness. Test specimens of materials that must meet the requirements 
of paragraph (a)(1)(v) of part I of this appendix must be no more than 
\1/8\-inch in thickness. Electrical wire and cable specimens must be the 
same size as used in the airplane. In the case of fabrics, both the warp 
and fill direction of the weave must be tested to determine the most 
critical flammability condition. Specimens must be mounted in a metal 
frame so that the two long edges and the upper edge are held securely 
during the vertical test prescribed in subparagraph (4) of this 
paragraph and the two long edges and the edge away from the flame are 
held securely during the horizontal test prescribed in subparagraph (5) 
of this paragraph. The exposed area of the specimen must be at least 2 
inches wide and 12 inches long, unless the actual size used in the 
airplane is smaller. The edge to which the burner flame is applied must 
not consist of the finished or protected edge of the specimen but must 
be representative of the actual cross-section of the material or part as 
installed in the airplane. The specimen must be mounted in a metal frame 
so that all four edges are held securely and the exposed area of the 
specimen is at least 8 inches by 8 inches during the 45[deg] test 
prescribed in subparagraph (6) of this paragraph.
    (3) Apparatus. Except as provided in subparagraph (7) of this 
paragraph, tests must be conducted in a draft-free cabinet in accordance 
with Federal Test Method Standard 191 Model 5903 (revised Method 5902) 
for the vertical test, or Method 5906 for horizontal test (available 
from the General Services Administration, Business Service Center, 
Region 3, Seventh & D Streets SW., Washington, DC 20407). Specimens 
which are too large for the cabinet must be tested in similar draft-free 
conditions.
    (4) Vertical test. A minimum of three specimens must be tested and 
results averaged. For fabrics, the direction of weave corresponding to 
the most critical flammability conditions must be parallel to the 
longest dimension. Each specimen must be supported vertically. The 
specimen must be exposed to a Bunsen or Tirrill burner with a nominal 
\3/8\-inch I.D. tube adjusted to give a flame of 1\1/2\ inches in 
height. The minimum flame temperature measured by a calibrated 
thermocouple pyrometer in the center of the flame must be 1550 [deg]F. 
The lower edge of the specimen must be \3/4\-inch above the top edge of 
the burner. The flame must be applied to the center line of the lower 
edge of the specimen. For materials covered by paragraph (a)(1)(i) of 
part I of this appendix, the flame must be applied for 60 seconds and 
then removed. For materials covered by paragraph (a)(1)(ii) of part I of 
this appendix, the flame must be applied for 12 seconds and then 
removed. Flame time, burn length, and flaming time of drippings, if any, 
may be recorded. The burn

[[Page 564]]

length determined in accordance with subparagraph (7) of this paragraph 
must be measured to the nearest tenth of an inch.
    (5) Horizontal test. A minimum of three specimens must be tested and 
the results averaged. Each specimen must be supported horizontally. The 
exposed surface, when installed in the aircraft, must be face down for 
the test. The specimen must be exposed to a Bunsen or Tirrill burner 
with a nominal \3/8\-inch I.D. tube adjusted to give a flame of 1\1/2\ 
inches in height. The minimum flame temperature measured by a calibrated 
thermocouple pyrometer in the center of the flame must be 1550 [deg]F. 
The specimen must be positioned so that the edge being tested is 
centered \3/4\-inch above the top of the burner. The flame must be 
applied for 15 seconds and then removed. A minimum of 10 inches of 
specimen must be used for timing purposes, approximately 1\1/2\ inches 
must burn before the burning front reaches the timing zone, and the 
average burn rate must be recorded.
    (6) Forty-five degree test. A minimum of three specimens must be 
tested and the results averaged. The specimens must be supported at an 
angle of 45[deg] to a horizontal surface. The exposed surface when 
installed in the aircraft must be face down for the test. The specimens 
must be exposed to a Bunsen or Tirrill burner with a nominal \3/8\-inch 
I.D. tube adjusted to give a flame of 1\1/2\ inches in height. The 
minimum flame temperature measured by a calibrated thermocouple 
pyrometer in the center of the flame must be 1550 [deg]F. Suitable 
precautions must be taken to avoid drafts. The flame must be applied for 
30 seconds with one-third contacting the material at the center of the 
specimen and then removed. Flame time, glow time, and whether the flame 
penetrates (passes through) the specimen must be recorded.
    (7) Sixty degree test. A minimum of three specimens of each wire 
specification (make and size) must be tested. The specimen of wire or 
cable (including insulation) must be placed at an angle of 60[deg] with 
the horizontal in the cabinet specified in subparagraph (3) of this 
paragraph with the cabinet door open during the test, or must be placed 
within a chamber approximately 2 feet high by 1 foot by 1 foot, open at 
the top and at one vertical side (front), and which allows sufficient 
flow of air for complete combustion, but which is free from drafts. The 
specimen must be parallel to and approximately 6 inches from the front 
of the chamber. The lower end of the specimen must be held rigidly 
clamped. The upper end of the specimen must pass over a pulley or rod 
and must have an appropriate weight attached to it so that the specimen 
is held tautly throughout the flammability test. The test specimen span 
between lower clamp and upper pulley or rod must be 24 inches and must 
be marked 8 inches from the lower end to indicate the central point for 
flame application. A flame from a Bunsen or Tirrill burner must be 
applied for 30 seconds at the test mark. The burner must be mounted 
underneath the test mark on the specimen, perpendicular to the specimen 
and at an angle of 30[deg] to the vertical plane of the specimen. The 
burner must have a nominal bore of \3/8\-inch and be adjusted to provide 
a 3-inch high flame with an inner cone approximately one-third of the 
flame height. The minimum temperature of the hottest portion of the 
flame, as measured with a calibrated thermocouple pyrometer, may not be 
less than 1750 [deg]F. The burner must be positioned so that the hottest 
portion of the flame is applied to the test mark on the wire. Flame 
time, burn length, and flaming time of drippings, if any, must be 
recorded. The burn length determined in accordance with paragraph (8) of 
this paragraph must be measured to the nearest tenth of an inch. 
Breaking of the wire specimens is not considered a failure.
    (8) Burn length. Burn length is the distance from the original edge 
to the farthest evidence of damage to the test specimen due to flame 
impingement, including areas of partial or complete consumption, 
charring, or embrittlement, but not including areas sooted, stained, 
warped, or discolored, nor areas where material has shrunk or melted 
away from the heat source.

                 Part II--Flammability of Seat Cushions

    (a) Criteria for Acceptance. Each seat cushion must meet the 
following criteria:
    (1) At least three sets of seat bottom and seat back cushion 
specimens must be tested.
    (2) If the cushion is constructed with a fire blocking material, the 
fire blocking material must completely enclose the cushion foam core 
material.
    (3) Each specimen tested must be fabricated using the principal 
components (i.e., foam core, flotation material, fire blocking material, 
if used, and dress covering) and assembly processes (representative 
seams and closures) intended for use in the production articles. If a 
different material combination is used for the back cushion than for the 
bottom cushion, both material combinations must be tested as complete 
specimen sets, each set consisting of a back cushion specimen and a 
bottom cushion specimen. If a cushion, including outer dress covering, 
is demonstrated to meet the requirements of this appendix using the oil 
burner test, the dress covering of that cushion may be replaced with a 
similar dress covering provided the burn length of the replacement 
covering, as determined by the test specified in Sec. 25.853(c), does 
not exceed the corresponding burn length of the dress covering used on 
the cushion subjected to the oil burner test.
    (4) For at least two-thirds of the total number of specimen sets 
tested, the burn length from the burner must not reach the side of the 
cushion opposite the burner. The

[[Page 565]]

burn length must not exceed 17 inches. Burn length is the perpendicular 
distance from the inside edge of the seat frame closest to the burner to 
the farthest evidence of damage to the test specimen due to flame 
impingement, including areas of partial or complete consumption, 
charring, or embrittlement, but not including areas sooted, stained, 
warped, or discolored, or areas where material has shrunk or melted away 
from the heat source.
    (5) The average percentage weight loss must not exceed 10 percent. 
Also, at least two-thirds of the total number of specimen sets tested 
must not exceed 10 percent weight loss. All droppings falling from the 
cushions and mounting stand are to be discarded before the after-test 
weight is determined. The percentage weight loss for a specimen set is 
the weight of the specimen set before testing less the weight of the 
specimen set after testing expressed as the percentage of the weight 
before testing.
    (b) Test Conditions. Vertical air velocity should average 25 
fpm10 fpm at the top of the back seat cushion. 
Horizontal air velocity should be below 10 fpm just above the bottom 
seat cushion. Air velocities should be measured with the ventilation 
hood operating and the burner motor off.
    (c) Test Specimens. (1) For each test, one set of cushion specimens 
representing a seat bottom and seat back cushion must be used.
    (2) The seat bottom cushion specimen must be 18 \1/8\ inches (457 3 mm) wide by 20 
\1/8\ inches (508 3 mm) deep 
by 4 \1/8\ inches (102 3 mm) 
thick, exclusive of fabric closures and seam overlap.
    (3) The seat back cushion specimen must be 18 \1/8\ inches (432 3 mm) wide by 25 
\1/8\ inches (635 3 mm) high 
by 2 \1/8\ inches (51 3 mm) 
thick, exclusive of fabric closures and seam overlap.
    (4) The specimens must be conditioned at 70 5 
[deg]F (21 2 [deg]C) 55%10% 
relative humidity for at least 24 hours before testing.
    (d) Test Apparatus. The arrangement of the test apparatus is shown 
in Figures 1 through 5 and must include the components described in this 
section. Minor details of the apparatus may vary, depending on the model 
burner used.
    (1) Specimen Mounting Stand. The mounting stand for the test 
specimens consists of steel angles, as shown in Figure 1. The length of 
the mounting stand legs is 12 \1/8\ inches (305 
3 mm). The mounting stand must be used for 
mounting the test specimen seat bottom and seat back, as shown in Figure 
2. The mounting stand should also include a suitable drip pan lined with 
aluminum foil, dull side up.
    (2) Test Burner. The burner to be used in testing must--
    (i) Be a modified gun type;
    (ii) Have an 80-degree spray angle nozzle nominally rated for 2.25 
gallons/hour at 100 psi;
    (iii) Have a 12-inch (305 mm) burner cone installed at the end of 
the draft tube, with an opening 6 inches (152 mm) high and 11 inches 
(280 mm) wide, as shown in Figure 3; and
    (iv) Have a burner fuel pressure regulator that is adjusted to 
deliver a nominal 2.0 gallon/hour of  2 Grade kerosene or equivalent 
required for the test.

Burner models which have been used successfully in testing are the 
Lennox Model OB-32, Carlin Model 200 CRD, and Park Model DPL 3400. FAA 
published reports pertinent to this type of burner are: (1) Powerplant 
Enginering Report No. 3A, Standard Fire Test Apparatus and Procedure for 
Flexible Hose Assemblies, dated March 1978; and (2) Report No. DOT/FAA/
RD/76/213, Reevaluation of Burner Characteristics for Fire Resistance 
Tests, dated January 1977.
    (3) Calorimeter.
    (i) The calorimeter to be used in testing must be a (0-15.0 BTU/
ft\2\-sec. 0-17.0 W/cm\2\) calorimeter, accurate 3%, mounted in a 6-inch by 12-inch (152 by 305 mm) by 
\3/4\-inch (19 mm) thick calcium silicate insulating board which is 
attached to a steel angle bracket for placement in the test stand during 
burner calibration, as shown in Figure 4.
    (ii) Because crumbling of the insulating board with service can 
result in misalignment of the calorimeter, the calorimeter must be 
monitored and the mounting shimmed, as necessary, to ensure that the 
calorimeter face is flush with the exposed plane of the insulating board 
in a plane parallel to the exit of the test burner cone.
    (4) Thermocouples. The seven thermocouples to be used for testing 
must be \1/16\- to \1/8\-inch metal sheathed, ceramic packed, type K, 
grounded thermocouples with a nominal 22 to 30 American wire gage (AWG)-
size conductor. The seven thermocouples must be attached to a steel 
angle bracket to form a thermocouple rake for placement in the test 
stand during burner calibration, as shown in Figure 5.
    (5) Apparatus Arrangement. The test burner must be mounted on a 
suitable stand to position the exit of the burner cone a distance of 4 
\1/8\ inches (102 3 mm) from 
one side of the specimen mounting stand. The burner stand should have 
the capability of allowing the burner to be swung away from the specimen 
mounting stand during warmup periods.
    (6) Data Recording. A recording potentiometer or other suitable 
calibrated instrument with an appropriate range must be used to measure 
and record the outputs of the calorimeter and the thermocouples.
    (7) Weight Scale. Weighing Device--A device must be used that with 
proper procedures may determine the before and after test weights of 
each set of seat cushion specimens within 0.02 pound (9 grams). A 
continuous weighing system is preferred.

[[Page 566]]

    (8) Timing Device. A stopwatch or other device (calibrated to 1 second) must be used to measure the time of 
application of the burner flame and self-extinguishing time or test 
duration.
    (e) Preparation of Apparatus. Before calibration, all equipment must 
be turned on and the burner fuel must be adjusted as specified in 
paragraph (d)(2).
    (f) Calibration. To ensure the proper thermal output of the burner, 
the following test must be made:
    (1) Place the calorimeter on the test stand as shown in Figure 4 at 
a distance of 4 \1/8\ inches (102 3 mm) from the exit of the burner cone.
    (2) Turn on the burner, allow it to run for 2 minutes for warmup, 
and adjust the burner air intake damper to produce a reading of 10.5 
0.5 BTU/ft\2\-sec. (11.9 0.6 
w/cm\2\) on the calorimeter to ensure steady state conditions have been 
achieved. Turn off the burner.
    (3) Replace the calorimeter with the thermocouple rake (Figure 5).
    (4) Turn on the burner and ensure that the thermocouples are reading 
1900 100 [deg]F (1038 38 
[deg]C) to ensure steady state conditions have been achieved.
    (5) If the calorimeter and thermocouples do not read within range, 
repeat steps in paragraphs 1 through 4 and adjust the burner air intake 
damper until the proper readings are obtained. The thermocouple rake and 
the calorimeter should be used frequently to maintain and record 
calibrated test parameters. Until the specific apparatus has 
demonstrated consistency, each test should be calibrated. After 
consistency has been confirmed, several tests may be conducted with the 
pre-test calibration before and a calibration check after the series.
    (g) Test Procedure. The flammability of each set of specimens must 
be tested as follows:
    (1) Record the weight of each set of seat bottom and seat back 
cushion specimens to be tested to the nearest 0.02 pound (9 grams).
    (2) Mount the seat bottom and seat back cushion test specimens on 
the test stand as shown in Figure 2, securing the seat back cushion 
specimen to the test stand at the top.
    (3) Swing the burner into position and ensure that the distance from 
the exit of the burner cone to the side of the seat bottom cushion 
specimen is 4 \1/8\ inches (102 3 mm).
    (4) Swing the burner away from the test position. Turn on the burner 
and allow it to run for 2 minutes to provide adequate warmup of the 
burner cone and flame stabilization.
    (5) To begin the test, swing the burner into the test position and 
simultaneously start the timing device.
    (6) Expose the seat bottom cushion specimen to the burner flame for 
2 minutes and then turn off the burner. Immediately swing the burner 
away from the test position. Terminate test 7 minutes after initiating 
cushion exposure to the flame by use of a gaseous extinguishing agent 
(i.e., Halon or CO2).
    (7) Determine the weight of the remains of the seat cushion specimen 
set left on the mounting stand to the nearest 0.02 pound (9 grams) 
excluding all droppings.
    (h) Test Report. With respect to all specimen sets tested for a 
particular seat cushion for which testing of compliance is performed, 
the following information must be recorded:
    (1) An identification and description of the specimens being tested.
    (2) The number of specimen sets tested.
    (3) The initial weight and residual weight of each set, the 
calculated percentage weight loss of each set, and the calculated 
average percentage weight loss for the total number of sets tested.
    (4) The burn length for each set tested.

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Part III--Test Method To Determine Flame Penetration Resistance of Cargo 
                           Compartment Liners.

    (a) Criteria for Acceptance. (1) At least three specimens of cargo 
compartment sidewall or ceiling liner panels must be tested.
    (2) Each specimen tested must simulate the cargo compartment 
sidewall or ceiling liner panel, including any design features, such as 
joints, lamp assemblies, etc., the failure of which would affect the 
capability of the liner to safely contain a fire.
    (3) There must be no flame penetration of any specimen within 5 
minutes after application of the flame source, and the peak temperature 
measured at 4 inches above the upper surface of the horizontal test 
sample must not exceed 400 [deg]F.

[[Page 572]]

    (b) Summary of Method. This method provides a laboratory test 
procedure for measuring the capability of cargo compartment lining 
materials to resist flame penetration with a 2 gallon per hour (GPH) 2 
Grade kerosene or equivalent burner fire source. Ceiling and sidewall 
liner panels may be tested individually provided a baffle is used to 
simulate the missing panel. Any specimen that passes the test as a 
ceiling liner panel may be used as a sidewall liner panel.
    (c) Test Specimens. (1) The specimen to be tested must measure 16 
\1/8\ inches (406 3 mm) by 
24+\1/8\ inches (610 3 mm).
    (2) The specimens must be conditioned at 70 [deg]F.5 [deg]F. (21 [deg]C. 2 [deg]C.) 
and 55%5% humidity for at least 24 hours before 
testing.
    (d) Test Apparatus. The arrangement of the test apparatus, which is 
shown in Figure 3 of Part II and Figures 1 through 3 of this part of 
appendix F, must include the components described in this section. Minor 
details of the apparatus may vary, depending on the model of the burner 
used.
    (1) Specimen Mounting Stand. The mounting stand for the test 
specimens consists of steel angles as shown in Figure 1.
    (2) Test Burner. The burner to be used in tesing must--
    (i) Be a modified gun type.
    (ii) Use a suitable nozzle and maintain fuel pressure to yield a 2 
GPH fuel flow. For example: an 80 degree nozzle nominally rated at 2.25 
GPH and operated at 85 pounds per square inch (PSI) gage to deliver 2.03 
GPH.
    (iii) Have a 12 inch (305 mm) burner extension installed at the end 
of the draft tube with an opening 6 inches (152 mm) high and 11 inches 
(280 mm) wide as shown in Figure 3 of Part II of this appendix.
    (iv) Have a burner fuel pressure regulator that is adjusted to 
deliver a nominal 2.0 GPH of 2 Grade kerosene or equivalent.

Burner models which have been used successfully in testing are the Lenox 
Model OB-32, Carlin Model 200 CRD and Park Model DPL. The basic burner 
is described in FAA Powerplant Engineering Report No. 3A, Standard Fire 
Test Apparatus and Procedure for Flexible Hose Assemblies, dated March 
1978; however, the test settings specified in this appendix differ in 
some instances from those specified in the report.
    (3) Calorimeter. (i) The calorimeter to be used in testing must be a 
total heat flux Foil Type Gardon Gage of an appropriate range 
(approximately 0 to 15.0 British thermal unit (BTU) per ft.\2\ sec., 0-
17.0 watts/cm\2\). The calorimeter must be mounted in a 6 inch by 12 
inch (152 by 305 mm) by \3/4\ inch (19 mm) thick insulating block which 
is attached to a steel angle bracket for placement in the test stand 
during burner calibration as shown in Figure 2 of this part of this 
appendix.
    (ii) The insulating block must be monitored for deterioration and 
the mounting shimmed as necessary to ensure that the calorimeter face is 
parallel to the exit plane of the test burner cone.
    (4) Thermocouples. The seven thermocouples to be used for testing 
must be \1/16\ inch ceramic sheathed, type K, grounded thermocouples 
with a nominal 30 American wire gage (AWG) size conductor. The seven 
thermocouples must be attached to a steel angle bracket to form a 
thermocouple rake for placement in the test stand during burner 
calibration as shown in Figure 3 of this part of this appendix.
    (5) Apparatus Arrangement. The test burner must be mounted on a 
suitable stand to position the exit of the burner cone a distance of 8 
inches from the ceiling liner panel and 2 inches from the sidewall liner 
panel. The burner stand should have the capability of allowing the 
burner to be swung away from the test specimen during warm-up periods.
    (6) Instrumentation. A recording potentiometer or other suitable 
instrument with an appropriate range must be used to measure and record 
the outputs of the calorimeter and the thermocouples.
    (7) Timing Device. A stopwatch or other device must be used to 
measure the time of flame application and the time of flame penetration, 
if it occurs.
    (e) Preparation of Apparatus. Before calibration, all equipment must 
be turned on and allowed to stabilize, and the burner fuel flow must be 
adjusted as specified in paragraph (d)(2).
    (f) Calibration. To ensure the proper thermal output of the burner 
the following test must be made:
    (1) Remove the burner extension from the end of the draft tube. Turn 
on the blower portion of the burner without turning the fuel or igniters 
on. Measure the air velocity using a hot wire anemometer in the center 
of the draft tube across the face of the opening. Adjust the damper such 
that the air velocity is in the range of 1550 to 1800 ft./min. If tabs 
are being used at the exit of the draft tube, they must be removed prior 
to this measurement. Reinstall the draft tube extension cone.
    (2) Place the calorimeter on the test stand as shown in Figure 2 at 
a distance of 8 inches (203 mm) from the exit of the burner cone to 
simulate the position of the horizontal test specimen.
    (3) Turn on the burner, allow it to run for 2 minutes for warm-up, 
and adjust the damper to produce a calorimeter reading of 8.0 0.5 BTU per ft.\2\ sec. (9.1 0.6 
Watts/cm\2\).
    (4) Replace the calorimeter with the thermocouple rake (see Figure 
3).
    (5) Turn on the burner and ensure that each of the seven 
thermocouples reads 1700 [deg]F. 100 [deg]F. (927 
[deg]C. 38 [deg]C.) to ensure steady state 
conditions have been achieved. If the temperature is out of this range, 
repeat steps 2 through 5 until proper readings are obtained.

[[Page 573]]

    (6) Turn off the burner and remove the thermocouple rake.
    (7) Repeat (1) to ensure that the burner is in the correct range.
    (g) Test Procedure. (1) Mount a thermocouple of the same type as 
that used for calibration at a distance of 4 inches (102 mm) above the 
horizontal (ceiling) test specimen. The thermocouple should be centered 
over the burner cone.
    (2) Mount the test specimen on the test stand shown in Figure 1 in 
either the horizontal or vertical position. Mount the insulating 
material in the other position.
    (3) Position the burner so that flames will not impinge on the 
specimen, turn the burner on, and allow it to run for 2 minutes. Rotate 
the burner to apply the flame to the specimen and simultaneously start 
the timing device.
    (4) Expose the test specimen to the flame for 5 minutes and then 
turn off the burner. The test may be terminated earlier if flame 
penetration is observed.
    (5) When testing ceiling liner panels, record the peak temperature 
measured 4 inches above the sample.
    (6) Record the time at which flame penetration occurs if applicable.
    (h) Test Report. The test report must include the following:
    (1) A complete description of the materials tested including type, 
manufacturer, thickness, and other appropriate data.
    (2) Observations of the behavior of the test specimens during flame 
exposure such as delamination, resin ignition, smoke, ect., including 
the time of such occurrence.
    (3) The time at which flame penetration occurs, if applicable, for 
each of the three specimens tested.
    (4) Panel orientation (ceiling or sidewall).

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   Part IV--Test Method To Determine the Heat Release Rate From Cabin 
                   Materials Exposed to Radiant Heat.

    (a) Summary of Method. Three or more specimens representing the 
completed aircraft component are tested. Each test specimen is injected 
into an environmental chamber through which a constant flow of air 
passes. The specimen's exposure is determined by a radiant heat source 
adjusted to produce, on the specimen, the desired total heat flux of 3.5 
W/cm\2\. The specimen is tested with the exposed surface vertical. 
Combustion is initiated by piloted ignition. The combustion products 
leaving the chamber are monitored in order to calculate the release rate 
of heat.
    (b) Apparatus. The Ohio State University (OSU) rate of heat release 
apparatus, as described below, is used. This is a modified version of 
the rate of heat release apparatus standardized by the American Society 
of Testing and Materials (ASTM), ASTM E-906.
    (1) This apparatus is shown in Figures 1A and 1B of this part IV. 
All exterior surfaces of the apparatus, except the holding chamber, must 
be insulated with 1 inch (25 mm) thick, low density, high temperature, 
fiberglass board insulation. A gasketed door, through which the sample 
injection rod slides, must be used to form an airtight closure on the 
specimen hold chamber.
    (2) Thermopile. The temperature difference between the air entering 
the environmental chamber and that leaving must be monitored by a 
thermopile having five hot, and five cold, 24-guage Chromel-Alumel 
junctions. The hot junctions must be spaced across the top of the 
exhaust stack, .38 inches (10 mm) below the top of the chimney. The 
thermocouples must have a .050 .010 inch (1.3 
.3mm) diameter, ball-type, welded tip. One 
thermocouple must be located in the geometric center, with the other 
four located 1.18 inch (30 mm) from the center along the diagonal toward 
each of the corners (Figure 5 of this part IV). The cold junctions must 
be located in the pan below the lower air distribution plate (see 
paragraph (b)(4) of this part IV). Thermopile hot junctions must be 
cleared of soot deposits as needed to maintain the calibrated 
sensitivity.
    (3) Radiation Source. A radiant heat source incorporating four Type 
LL silicon carbide elements, 20 inches (508 mm) long by .63 inch (16 mm) 
O.D., must be used, as shown in Figures 2A and 2B of this part IV. The 
heat source must have a nominal resistance of 1.4 ohms and be capable of 
generating a flux up to 100 kW/m\2\. The silicone carbide elements must 
be mounted in the stainless steel panel box by inserting them through 
.63 inch (16 mm) holes in .03 inch (1 mm) thick ceramic fiber or 
calcium-silicate millboard. Locations of the holes in the pads and 
stainless steel cover plates are shown in Figure 2B of this part IV. The 
truncated diamond-shaped mask of .042 .002 inch 
(1.07 .05mm) stainless steel must be added to 
provide uniform heat flux density over the area occupied by the vertical 
sample.
    (4) Air Distribution System. The air entering the environmental 
chamber must be distributed by a .25 inch (6.3 mm) thick aluminum plate 
having eight No. 4 drill-holes, located 2 inches (51 mm) from sides on 4 
inch (102 mm) centers, mounted at the base of the environmental chamber. 
A second plate of 18 guage stainless steel having 120, evenly spaced, 
No. 28 drill holes must be mounted 6 inches (152 mm) above the aluminum 
plate. A well-regulated air supply is required. The air-supply manifold 
at the base of the pyramidal section must have 48, evenly spaced, No. 26 
drill holes located .38 inch (10 mm) from the inner edge of the 
manifold, resulting in an airflow split of approximately three to one 
within the apparatus.
    (5) Exhaust Stack. An exhaust stack, 5.25x2.75 inches (133x70 mm) in 
cross section, and 10 inches (254 mm) long, fabricated from 28 guage 
stainless steel must be mounted on the outlet of the pyramidal section. 
A. 1.0x3.0 inch (25x76 mm) baffle plate of .018 .002 inch (.50 .05 mm) stainless 
steel must be centered inside the stack, perpendicular to the air flow, 
3 inches (76 mm) above the base of the stack.
    (6) Specimen Holders. (i) The specimen must be tested in a vertical 
orientation. The specimen holder (Figure 3 of this part IV) must 
incorporate a frame that touches the specimen (which is wrapped with 
aluminum foil as required by paragraph (d)(3) of this Part) along only 
the .25 inch (6 mm) perimeter. A ``V'' shaped spring is used to hold the 
assembly together. A detachable .50x50x5.91 inch (12x12x150 mm) drip pan 
and two .020 inch (.5 mm) stainless steel wires (as shown in Figure 3 of 
this part IV) must be used for testing materials prone to melting and 
dripping. The positioning of the spring and frame may be changed to 
accommodate different specimen thicknesses by inserting the retaining 
rod in different holes on the specimen holder.
    (ii) Since the radiation shield described in ASTM E-906 is not used, 
a guide pin must be added to the injection mechanism. This fits into a 
slotted metal plate on the injection mechanism outside of the holding 
chamber. It can be used to provide accurate positioning of the specimen 
face after injection. The front surface of the specimen must be 3.9 
inches (100 mm) from the closed radiation doors after injection.
    (iii) The specimen holder clips onto the mounted bracket (Figure 3 
of this part IV). The mounting bracket must be attached to the injection 
rod by three screws that pass through a wide-area washer welded onto a 
\1/2\-inch (13 mm) nut. The end of the injection rod must be threaded to 
screw into the nut, and a .020 inch (5.1 mm) thick wide area

[[Page 578]]

washer must be held between two \1/2\-inch (13 mm) nuts that are 
adjusted to tightly cover the hole in the radiation doors through which 
the injection rod or calibration calorimeter pass.
    (7) Calorimeter. A total-flux type calorimeter must be mounted in 
the center of a \1/2\-inch Kaowool ``M'' board inserted in the sample 
holder to measure the total heat flux. The calorimeter must have a view 
angle of 180 degrees and be calibrated for incident flux. The 
calorimeter calibration must be acceptable to the Administrator.
    (8) Pilot-Flame Positions. Pilot ignition of the specimen must be 
accomplished by simultaneously exposing the specimen to a lower pilot 
burner and an upper pilot burner, as described in paragraph (b)(8)(i) 
and (b)(8)(ii) or (b)(8)(iii) of this part IV, respectively. Since 
intermittent pilot flame extinguishment for more than 3 seconds would 
invalidate the test results, a spark ignitor may be installed to ensure 
that the lower pilot burner remains lighted.
    (i) Lower Pilot Burner. The pilot-flame tubing must be .25 inch (6.3 
mm) O.D., .03 inch (0.8mm) wall, stainless steel tubing. A mixture of 
120 cm\3\/min. of methane and 850 cm\3\/min. of air must be fed to the 
lower pilot flame burner. The normal position of the end of the pilot 
burner tubing is .40 inch (10 mm) from and perpendicular to the exposed 
vertical surface of the specimen. The centerline at the outlet of the 
burner tubing must intersect the vertical centerline of the sample at a 
point .20 inch (5 mm) above the lower exposed edge of the specimen.
    (ii) Standard Three-Hole Upper Pilot Burner. The pilot burner must 
be a straight length of .25 inch (6.3 mm) O.D., .03 inch (0.8 mm) wall, 
stainless steel tubing that is 14 inches (360 mm) long. One end of the 
tubing must be closed, and three No. 40 drill holes must be drilled into 
the tubing, 2.38 inch (60 mm) apart, for gas ports, all radiating in the 
same direction. The first hole must be .19 inch (5 mm) from the closed 
end of the tubing. The tube must be positioned .75 inch (19 mm) above 
and .75 inch (19 mm) behind the exposed upper edge of the specimen. The 
middle hole must be in the vertical plane perpendicular to the exposed 
surface of the specimen which passes through its vertical centerline and 
must be pointed toward the radiation source. The gas supplied to the 
burner must be methane and must be adjusted to produce flame lengths of 
1 inch (25 mm).
    (iii) Optional Fourteen-Hole Upper Pilot Burner. This burner may be 
used in lieu of the standard three-hole burner described in paragraph 
(b)(8)(ii) of this part IV. The pilot burner must be a straight length 
of .25 inch (6.3 mm) O.D., .03 inch (0.8 mm) wall, stainless steel 
tubing that is 15.75 inches (400 mm) long. One end of the tubing must be 
closed, and 14 No. 59 drill holes must be drilled into the tubing, .50 
inch (13 mm) apart, for gas ports, all radiating in the same direction. 
The first hole must be .50 inch (13 mm) from the closed end of the 
tubing. The tube must be positioned above the specimen holder so that 
the holes are placed above the specimen as shown in Figure 1B of this 
part IV. The fuel supplied to the burner must be methane mixed with air 
in a ratio of approximately 50/50 by volume. The total gas flow must be 
adjusted to produce flame lengths of 1 inch (25 mm). When the gas/air 
ratio and the flow rate are properly adjusted, approximately .25 inch (6 
mm) of the flame length appears yellow in color.
    (c) Calibration of Equipment--(1) Heat Release Rate. A calibration 
burner, as shown in Figure 4, must be placed over the end of the lower 
pilot flame tubing using a gas tight connection. The flow of gas to the 
pilot flame must be at least 99 percent methane and must be accurately 
metered. Prior to usage, the wet test meter must be properly leveled and 
filled with distilled water to the tip of the internal pointer while no 
gas is flowing. Ambient temperature and pressure of the water are based 
on the internal wet test meter temperature. A baseline flow rate of 
approximately 1 liter/min. must be set and increased to higher preset 
flows of 4, 6, 8, 6 and 4 liters/min. Immediately prior to recording 
methane flow rates, a flow rate of 8 liters/min. must be used for 2 
minutes to precondition the chamber. This is not recorded as part of 
calibration. The rate must be determined by using a stopwatch to time a 
complete revolution of the wet test meter for both the baseline and 
higher flow, with the flow returned to baseline before changing to the 
next higher flow. The thermopile baseline voltage must be measured. The 
gas flow to the burner must be increased to the higher preset flow and 
allowed to burn for 2.0 minutes, and the thermopile voltage must be 
measured. The sequence must be repeated until all five values have been 
determined. The average of the five values must be used as the 
calibration factor. The procedure must be repeated if the percent 
relative standard deviation is greater than 5 percent. Calculations are 
shown in paragraph (f) of this part IV.
    (2) Flux Uniformity. Uniformity of flux over the specimen must be 
checked periodically and after each heating element change to determine 
if it is within acceptable limits of plus or minus 5 percent.
    (3) As noted in paragraph (b)(2) of this part IV, thermopile hot 
junctions must be cleared of soot deposits as needed to maintain the 
calibrated sensitivity.
    (d) Preparation of Test Specimens. (1) The test specimens must be 
representative of the aircraft component in regard to materials and 
construction methods. The standard size for the test specimens is 5.91 
.03 x 5.91 .03 inches (149 
1 x 149 1 mm). The thickness 
of the specimen must be the same as that of the

[[Page 579]]

aircraft component it represents up to a maximum thickness of 1.75 
inches (45 mm). Test specimens representing thicker components must be 
1.75 inches (45 mm).
    (2) Conditioning. Specimens must be conditioned as described in Part 
1 of this appendix.
    (3) Mounting. Each test specimen must be wrapped tightly on all 
sides of the specimen, except for the one surface that is exposed with a 
single layer of .001 inch (.025 mm) aluminum foil.
    (e) Procedure. (1) The power supply to the radiant panel must be set 
to produce a radiant flux of 3.5 .05 W/cm\2\, as 
measured at the point the center of the specimen surface will occupy 
when positioned for the test. The radiant flux must be measured after 
the air flow through the equipment is adjusted to the desired rate.
    (2) After the pilot flames are lighted, their position must be 
checked as described in paragraph (b)(8) of this part IV.
    (3) Air flow through the apparatus must be controlled by a circular 
plate orifice located in a 1.5 inch (38.1 mm) I.D. pipe with two 
pressure measuring points, located 1.5 inches (38 mm) upstream and .75 
inches (19 mm) downstream of the orifice plate. The pipe must be 
connected to a manometer set at a pressure differential of 7.87 inches 
(200 mm) of Hg. (See Figure 1B of this part IV.) The total air flow to 
the equipment is approximately .04 m\3\/seconds. The stop on the 
vertical specimen holder rod must be adjusted so that the exposed 
surface of the specimen is positioned 3.9 inches (100 mm) from the 
entrance when injected into the environmental chamber.
    (4) The specimen must be placed in the hold chamber with the 
radiation doors closed. The airtight outer door must be secured, and the 
recording devices must be started. The specimen must be retained in the 
hold chamber for 60 seconds, plus or minus 10 seconds, before injection. 
The thermopile ``zero'' value must be determined during the last 20 
seconds of the hold period. The sample must not be injected before 
completion of the ``zero'' value determination.
    (5) When the specimen is to be injected, the radiation doors must be 
opened. After the specimen is injected into the environmental chamber, 
the radiation doors must be closed behind the specimen.
    (6) [Reserved]
    (7) Injection of the specimen and closure of the inner door marks 
time zero. A record of the thermopile output with at least one data 
point per second must be made during the time the specimen is in the 
environmental chamber.
    (8) The test duration is five minutes. The lower pilot burner and 
the upper pilot burner must remain lighted for the entire duration of 
the test, except that there may be intermittent flame extinguishment for 
periods that do not exceed 3 seconds. Furthermore, if the optional 
three-hole upper burner is used, at least two flamelets must remain 
lighted for the entire duration of the test, except that there may be 
intermittent flame extinguishment of all three flamelets for periods 
that do not exceed 3 seconds.
    (9) A minimum of three specimens must be tested.
    (f) Calculations. (1) The calibration factor is calculated as 
follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.072

F0=flow of methane at baseline (1pm)
F1=higher preset flow of methane (1pm)
V0=thermopile voltage at baseline (mv)
V1=thermopile voltage at higher flow (mv)
Ta=Ambient temperature (K)
P=Ambient pressure (mm Hg)
Pv=Water vapor pressure (mm Hg)

    (2) Heat release rates may be calculated from the reading of the 
thermopile output voltage at any instant of time as:
[GRAPHIC] [TIFF OMITTED] TR02FE95.006

HRR=heat release rate (kw/m\2\)
Vb=baseline voltage (mv)
Vm=measured thermopile voltage (mv)
Kh=calibration factor (kw/mv)

    (3) The integral of the heat release rate is the total heat release 
as a function of time and is calculated by multiplying the rate by the 
data sampling frequency in minutes and summing the time from zero to two 
minutes.
    (g) Criteria. The total positive heat release over the first two 
minutes of exposure for each of the three or more samples tested must be 
averaged, and the peak heat release rate for each of the samples must be 
averaged. The average total heat release must not exceed 65 kilowatt-
minutes per square meter, and the average peak heat release rate must 
not exceed 65 kilowatts per square meter.
    (h) Report. The test report must include the following for each 
specimen tested:

[[Page 580]]

    (1) Description of the specimen.
    (2) Radiant heat flux to the specimen, expressed in W/cm\2\.
    (3) Data giving release rates of heat (in kW/m\2\ ) as a function of 
time, either graphically or tabulated at intervals no greater than 10 
seconds. The calibration factor (kn) must be recorded.
    (4) If melting, sagging, delaminating, or other behavior that 
affects the exposed surface area or the mode of burning occurs, these 
behaviors must be reported, together with the time at which such 
behaviors were observed.
    (5) The peak heat release and the 2-minute integrated heat release 
rate must be reported.

[[Page 581]]

                    Figures to Part IV of Appendix F
[GRAPHIC] [TIFF OMITTED] TR02FE95.004


[[Page 582]]


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[[Page 583]]


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[[Page 584]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.074


[[Page 585]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.075


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[GRAPHIC] [TIFF OMITTED] TC28SE91.076


[[Page 587]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.077

 Part V. Test Method To Determine the Smoke Emission Characteristics of 
                             Cabin Materials

    (a) Summary of Method. The specimens must be constructed, 
conditioned, and tested in the flaming mode in accordance with American 
Society of Testing and Materials (ASTM) Standard Test Method ASTM F814-
83.
    (b) Acceptance Criteria. The specific optical smoke density 
(Ds), which is obtained by averaging the reading obtained 
after 4 minutes with each of the three specimens, shall not exceed 200.

Part VI--Test Method To Determine the Flammability and Flame Propagation 
        Characteristics of Thermal/Acoustic Insulation Materials

    Use this test method to evaluate the flammability and flame 
propagation characteristics of thermal/acoustic insulation when exposed 
to both a radiant heat source and a flame.
    (a) Definitions.
    ``Flame propagation'' means the furthest distance of the propagation 
of visible flame towards the far end of the test specimen, measured from 
the midpoint of the ignition source flame. Measure this distance after 
initially applying the ignition source and before all flame on the test 
specimen is extinguished. The measurement is not a determination of burn 
length made after the test.

[[Page 588]]

    ``Radiant heat source'' means an electric or air propane panel.
    ``Thermal/acoustic insulation'' means a material or system of 
materials used to provide thermal and/or acoustic protection. Examples 
include fiberglass or other batting material encapsulated by a film 
covering and foams.
    ``Zero point'' means the point of application of the pilot burner to 
the test specimen.
    (b) Test apparatus.
    [GRAPHIC] [TIFF OMITTED] TR31JY03.003
    
    (1) Radiant panel test chamber. Conduct tests in a radiant panel 
test chamber (see figure 1 above). Place the test chamber under an 
exhaust hood to facilitate clearing the chamber of smoke after each 
test. The radiant panel test chamber must be an enclosure 55 inches 
(1397 mm) long by 19.5 (495 mm) deep by 28 (710 mm) to 30 inches 
(maximum) (762 mm) above the test specimen. Insulate the sides, ends, 
and top with a fibrous ceramic insulation, such as Kaowool M\TM\ board. 
On the front side, provide a 52 by 12-inch (1321 by 305 mm) draft-free, 
high-temperature, glass window for viewing the sample during testing. 
Place a door below the window to provide access to the movable specimen 
platform holder. The bottom of the test chamber must be a sliding steel 
platform that has provision for securing the test specimen holder in a 
fixed and level position. The chamber must have an internal chimney with 
exterior dimensions of 5.1 inches (129 mm) wide, by 16.2 inches (411 mm) 
deep by 13 inches (330 mm) high at the opposite end of the chamber from 
the radiant energy source. The interior dimensions must be 4.5 inches 
(114 mm) wide by 15.6 inches (395 mm) deep. The chimney must extend to 
the top of the chamber (see figure 2).

[[Page 589]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.004

    (2) Radiant heat source. Mount the radiant heat energy source in a 
cast iron frame or equivalent. An electric panel must have six, 3-inch 
wide emitter strips. The emitter strips must be perpendicular to the 
length of the panel. The panel must have a radiation surface of 12\7/8\ 
by 18\1/2\ inches (327 by 470 mm). The panel must be capable of 
operating at temperatures up to 1300 [deg]F (704 [deg]C). An air propane 
panel must be made of a porous refractory material and have a radiation 
surface of 12 by 18 inches (305 by 457 mm). The panel must be capable of 
operating at temperatures up to 1,500 [deg]F (816 [deg]C). See figures 
3a and 3b.

[[Page 590]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.005


[[Page 591]]


[GRAPHIC] [TIFF OMITTED] TR31JY03.006

    (i) Electric radiant panel. The radiant panel must be 3-phase and 
operate at 208 volts. A single-phase, 240 volt panel is also acceptable. 
Use a solid-state power controller and microprocessor-based controller 
to set the electric panel operating parameters.
    (ii) Gas radiant panel. Use propane (liquid petroleum gas--2.1 UN 
1075) for the radiant panel fuel. The panel fuel system must consist of 
a venturi-type aspirator for mixing gas and air at approximately 
atmospheric pressure. Provide suitable instrumentation for monitoring 
and controlling the flow of fuel and air to the panel. Include an air 
flow gauge, an air flow regulator, and a gas pressure gauge.
    (iii) Radiant panel placement. Mount the panel in the chamber at 
30[deg] to the horizontal specimen plane, and 7\1/2\ inches above the 
zero point of the specimen.
    (3) Specimen holding system.
    (i) The sliding platform serves as the housing for test specimen 
placement. Brackets may be attached (via wing nuts) to the top lip of 
the platform in order to accommodate various thicknesses of test 
specimens. Place the test specimens on a sheet of Kaowool M\TM\ board or 
1260 Standard Board (manufactured by Thermal Ceramics and available in 
Europe), or equivalent, either resting on the bottom lip of the sliding 
platform or on the base of the brackets. It may be necessary to use 
multiple sheets of material based on the thickness of the test specimen 
(to meet the sample height requirement). Typically, these non-
combustible sheets of material are available in \1/4\ inch (6 mm) 
thicknesses. See figure 4. A sliding platform that is deeper than the 2-
inch (50.8mm) platform shown in figure 4 is also acceptable as long as 
the sample height requirement is met.

[[Page 592]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.007

    (ii) Attach a \1/2\ inch (13 mm) piece of Kaowool M\TM\ board or 
other high temperature material measuring 41\1/2\ by 8\1/4\ inches (1054 
by 210 mm) to the back of the platform. This board serves as a heat 
retainer and protects the test specimen from excessive preheating. The 
height of this board must not impede the sliding platform movement (in 
and out of the test chamber). If the platform has been fabricated such 
that the back side of the platform is high enough to prevent excess 
preheating of the specimen when the sliding platform is out, a retainer 
board is not necessary.
    (iii) Place the test specimen horizontally on the non-combustible 
board(s). Place a steel retaining/securing frame fabricated of mild 
steel, having a thickness of \1/8\ inch (3.2 mm) and overall dimensions 
of 23 by 13\1/8\ inches (584 by 333 mm) with a specimen opening of 19 by 
10\3/4\ inches (483 by 273 mm) over the test specimen. The front, back, 
and right portions of the top flange of the frame must rest on the top 
of the sliding platform, and the bottom flanges must pinch all 4 sides 
of the test specimen. The right bottom flange must be flush with the 
sliding platform. See figure 5.

[[Page 593]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.008

    (4) Pilot Burner. The pilot burner used to ignite the specimen must 
be a Bernzomatic\TM\ commercial propane venturi torch with an axially 
symmetric burner tip and a propane supply tube with an orifice diameter 
of 0.006 inches (0.15 mm). The length of the burner tube must be 2\7/8\ 
inches (71 mm). The propane flow must be adjusted via gas pressure 
through an in-line regulator to produce a blue inner cone length of \3/
4\ inch (19 mm). A \3/4\ inch (19 mm) guide (such as a thin strip of 
metal) may be soldered to the top of the burner to aid in setting the 
flame height. The overall flame length must be approximately 5 inches 
long (127 mm). Provide a way to move the burner out of the ignition 
position so that the flame is horizontal and at least 2 inches (50 mm) 
above the specimen plane. See figure 6.

[[Page 594]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.009

    (5) Thermocouples. Install a 24 American Wire Gauge (AWG) Type K 
(Chromel-Alumel) thermocouple in the test chamber for temperature 
monitoring. Insert it into the chamber through a small hole drilled 
through the back of the chamber. Place the thermocouple so that it 
extends 11 inches (279 mm) out from the back of the chamber wall, 11\1/
2\ inches (292 mm) from the right side of the chamber wall, and is 2 
inches (51 mm) below the radiant panel. The use of other thermocouples 
is optional.
    (6) Calorimeter. The calorimeter must be a one-inch cylindrical 
water-cooled, total heat flux density, foil type Gardon Gage that has a 
range of 0 to 5 BTU/ft\2\-second (0 to 5.7 Watts/cm\2\).
    (7) Calorimeter calibration specification and procedure.
    (i) Calorimeter specification.
    (A) Foil diameter must be 0.25 0.005 inches 
(6.35 0.13 mm).
    (B) Foil thickness must be 0.0005 0.0001 
inches (0.013 0.0025 mm).
    (C) Foil material must be thermocouple grade Constantan.
    (D) Temperature measurement must be a Copper Constantan 
thermocouple.
    (E) The copper center wire diameter must be 0.0005 inches (0.013 
mm).
    (F) The entire face of the calorimeter must be lightly coated with 
``Black Velvet'' paint having an emissivity of 96 or greater.
    (ii) Calorimeter calibration. (A) The calibration method must be by 
comparison to a like standardized transducer.
    (B) The standardized transducer must meet the specifications given 
in paragraph VI(b)(6) of this appendix.
    (C) Calibrate the standard transducer against a primary standard 
traceable to the National Institute of Standards and Technology (NIST).
    (D) The method of transfer must be a heated graphite plate.
    (E) The graphite plate must be electrically heated, have a clear 
surface area on each side of the plate of at least 2 by 2 inches (51 by 
51 mm), and be \1/8\ inch \1/16\ inch thick (3.2 
1.6 mm).
    (F) Center the 2 transducers on opposite sides of the plates at 
equal distances from the plate.
    (G) The distance of the calorimeter to the plate must be no less 
than 0.0625 inches (1.6 mm), nor greater than 0.375 inches (9.5 mm).
    (H) The range used in calibration must be at least 0-3.5 BTUs/ft\2\ 
second (0-3.9 Watts/cm\2\) and no greater than 0-5.7 BTUs/ft\2\ second 
(0-6.4 Watts/cm\2\).
    (I) The recording device used must record the 2 transducers 
simultaneously or at least within \1/10\ of each other.
    (8) Calorimeter fixture. With the sliding platform pulled out of the 
chamber, install the calorimeter holding frame and place a sheet of non-
combustible material in the bottom of the sliding platform adjacent to 
the holding frame. This will prevent heat losses during calibration. The 
frame must be 13\1/8\ inches (333 mm) deep (front to back) by 8 inches 
(203 mm) wide and must rest on the top of the sliding platform. It must 
be fabricated of \1/8\ inch (3.2 mm) flat stock steel and have an 
opening that accommodates a \1/2\ inch (12.7 mm) thick piece of 
refractory board, which is level with the top of the sliding platform. 
The board must have three 1-inch (25.4 mm) diameter holes drilled 
through the board for calorimeter insertion. The distance to the radiant 
panel surface from the centerline of the first hole (``zero'' position) 
must be 7\1/2\ \1/8\ inches (191 3 mm). The distance between the centerline of the first 
hole to the centerline of the second hole must be 2 inches (51 mm). It 
must also be the same distance from the centerline of the second hole to 
the centerline of the third hole. See figure 7. A calorimeter holding 
frame that differs in construction is acceptable as long as the height 
from the centerline of the

[[Page 595]]

first hole to the radiant panel and the distance between holes is the 
same as described in this paragraph.
[GRAPHIC] [TIFF OMITTED] TR31JY03.010

    (9) Instrumentation. Provide a calibrated recording device with an 
appropriate range or a computerized data acquisition system to measure 
and record the outputs of the calorimeter and the thermocouple. The data 
acquisition system must be capable of recording the calorimeter output 
every second during calibration.
    (10) Timing device. Provide a stopwatch or other device, accurate to 
1 second/hour, to measure the time of application 
of the pilot burner flame.
    (c) Test specimens. (1) Specimen preparation. Prepare and test a 
minimum of three test specimens. If an oriented film cover material is 
used, prepare and test both the warp and fill directions.
    (2) Construction. Test specimens must include all materials used in 
construction of the insulation (including batting, film, scrim, tape 
etc.). Cut a piece of core material such as foam or fiberglass, and cut 
a piece of film cover material (if used) large enough to cover the core 
material. Heat sealing is the preferred method of preparing fiberglass 
samples, since they can be made without compressing the fiberglass 
(``box sample''). Cover materials that are not heat sealable may be 
stapled, sewn, or taped as long as the cover material is over-cut enough 
to be drawn down the sides without compressing the core material. The 
fastening means should be as continuous as possible along the length of 
the seams. The specimen thickness must be of the same thickness as 
installed in the airplane.
    (3) Specimen Dimensions. To facilitate proper placement of specimens 
in the sliding platform housing, cut non-rigid core materials, such as 
fiberglass, 12\1/2\ inches (318mm) wide by 23 inches (584mm) long. Cut 
rigid materials, such as foam, 11\1/2\ \1/4\ 
inches (292 mm 6mm) wide by 23 inches (584mm) long 
in order to fit properly in the sliding platform housing and provide a 
flat, exposed surface equal to the opening in the housing.
    (d) Specimen conditioning. Condition the test specimens at 70 5 [deg]F (21 2 [deg]C) and 55% 
10% relative humidity, for a minimum of 24 hours 
prior to testing.
    (e) Apparatus Calibration. (1) With the sliding platform out of the 
chamber, install the calorimeter holding frame. Push the platform back 
into the chamber and insert the calorimeter into the first hole 
(``zero'' position). See figure 7. Close the bottom door located below 
the sliding platform. The distance from the centerline of the 
calorimeter

[[Page 596]]

to the radiant panel surface at this point must be 7.\1/2\ inches \1/8\ (191 mm 3). Prior to 
igniting the radiant panel, ensure that the calorimeter face is clean 
and that there is water running through the calorimeter.
    (2) Ignite the panel. Adjust the fuel/air mixture to achieve 1.5 
BTUs/ft\2\-second 5% (1.7 Watts/cm\2\ 5%) at the ``zero'' position. If using an electric 
panel, set the power controller to achieve the proper heat flux. Allow 
the unit to reach steady state (this may take up to 1 hour). The pilot 
burner must be off and in the down position during this time.
    (3) After steady-state conditions have been reached, move the 
calorimeter 2 inches (51 mm) from the ``zero'' position (first hole) to 
position 1 and record the heat flux. Move the calorimeter to position 2 
and record the heat flux. Allow enough time at each position for the 
calorimeter to stabilize. Table 1 depicts typical calibration values at 
the three positions.

                       Table 1--Calibration Table
------------------------------------------------------------------------
          Position               BTU's/ft\2\sec          Watts/cm\2\
------------------------------------------------------------------------
``Zero'' Position...........                   1.5                   1.7
Position 1..................        1.51-1.50-1.49        1.71-1.70-1.69
Position 2..................             1.43-1.44             1.62-1.63
------------------------------------------------------------------------

    (4) Open the bottom door, remove the calorimeter and holder fixture. 
Use caution as the fixture is very hot.
    (f) Test Procedure. (1) Ignite the pilot burner. Ensure that it is 
at least 2 inches (51 mm) above the top of the platform. The burner must 
not contact the specimen until the test begins.
    (2) Place the test specimen in the sliding platform holder. Ensure 
that the test sample surface is level with the top of the platform. At 
``zero'' point, the specimen surface must be 7\1/2\ inches \1/8\ inch (191 mm 3) below the 
radiant panel.
    (3) Place the retaining/securing frame over the test specimen. It 
may be necessary (due to compression) to adjust the sample (up or down) 
in order to maintain the distance from the sample to the radiant panel 
(7\1/2\ inches \1/8\ inch (191 mm3) at ``zero'' position). With film/fiberglass 
assemblies, it is critical to make a slit in the film cover to purge any 
air inside. This allows the operator to maintain the proper test 
specimen position (level with the top of the platform) and to allow 
ventilation of gases during testing. A longitudinal slit, approximately 
2 inches (51mm) in length, must be centered 3 inches \1/2\ inch (76mm13mm) from the 
left flange of the securing frame. A utility knife is acceptable for 
slitting the film cover.
    (4) Immediately push the sliding platform into the chamber and close 
the bottom door.
    (5) Bring the pilot burner flame into contact with the center of the 
specimen at the ``zero'' point and simultaneously start the timer. The 
pilot burner must be at a 27[deg] angle with the sample and be 
approximately \1/2\ inch (12 mm) above the sample. See figure 7. A stop, 
as shown in figure 8, allows the operator to position the burner 
correctly each time.

[[Page 597]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.011

    (6) Leave the burner in position for 15 seconds and then remove to a 
position at least 2 inches (51 mm) above the specimen.
    (g) Report. (1) Identify and describe the test specimen.
    (2) Report any shrinkage or melting of the test specimen.
    (3) Report the flame propagation distance. If this distance is less 
than 2 inches, report this as a pass (no measurement required).
    (4) Report the after-flame time.
    (h) Requirements. (1) There must be no flame propagation beyond 2 
inches (51 mm) to the left of the centerline of the pilot flame 
application.
    (2) The flame time after removal of the pilot burner may not exceed 
3 seconds on any specimen.

    Part VII--Test Method To Determine the Burnthrough Resistance of 
                  Thermal/Acoustic Insulation Materials

    Use the following test method to evaluate the burnthrough resistance 
characteristics of aircraft thermal/acoustic insulation materials when 
exposed to a high intensity open flame.

[[Page 598]]

    (a) Definitions.
    Burnthrough time means the time, in seconds, for the burner flame to 
penetrate the test specimen, and/or the time required for the heat flux 
to reach 2.0 Btu/ft\2\sec (2.27 W/cm\2\) on the inboard side, at a 
distance of 12 inches (30.5 cm) from the front surface of the insulation 
blanket test frame, whichever is sooner. The burnthrough time is 
measured at the inboard side of each of the insulation blanket 
specimens.
    Insulation blanket specimen means one of two specimens positioned in 
either side of the test rig, at an angle of 30[deg] with respect to 
vertical.
    Specimen set means two insulation blanket specimens. Both specimens 
must represent the same production insulation blanket construction and 
materials, proportioned to correspond to the specimen size.
    (b) Apparatus. (1) The arrangement of the test apparatus is shown in 
figures 1 and 2 and must include the capability of swinging the burner 
away from the test specimen during warm-up.

[[Page 599]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.012

    (2) Test burner. The test burner must be a modified gun-type such as 
the Park Model DPL 3400. Flame characteristics are highly dependent on 
actual burner setup. Parameters such as fuel pressure, nozzle depth, 
stator position, and intake airflow must be properly adjusted to achieve 
the correct flame output.

[[Page 600]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.013

    (i) Nozzle. A nozzle must maintain the fuel pressure to yield a 
nominal 6.0 gal/hr (0.378 L/min) fuel flow. A Monarch-manufactured 
80[deg] PL (hollow cone) nozzle nominally rated at 6.0 gal/hr at 100 lb/
in\2\ (0.71 MPa) delivers a proper spray pattern.
    (ii) Fuel Rail. The fuel rail must be adjusted to position the fuel 
nozzle at a depth of 0.3125 inch (8 mm) from the end plane of the exit 
stator, which must be mounted in the end of the draft tube.
    (iii) Internal Stator. The internal stator, located in the middle of 
the draft tube, must

[[Page 601]]

be positioned at a depth of 3.75 inches (95 mm) from the tip of the fuel 
nozzle. The stator must also be positioned such that the integral 
igniters are located at an angle midway between the 10 and 11 o'clock 
position, when viewed looking into the draft tube. Minor deviations to 
the igniter angle are acceptable if the temperature and heat flux 
requirements conform to the requirements of paragraph VII(e) of this 
appendix.
    (iv) Blower Fan. The cylindrical blower fan used to pump air through 
the burner must measure 5.25 inches (133 mm) in diameter by 3.5 inches 
(89 mm) in width.
    (v) Burner cone. Install a 12 +0.125-inch (305 3 mm) burner extension cone at the end of the draft 
tube. The cone must have an opening 6 0.125-inch 
(152 3 mm) high and 11 0.125-inch (280 3 mm) wide (see 
figure 3).
    (vi) Fuel. Use JP-8, Jet A, or their international equivalent, at a 
flow rate of 6.0 0.2 gal/hr (0.378 0.0126 L/min). If this fuel is unavailable, ASTM K2 fuel 
(Number 2 grade kerosene) or ASTM D2 fuel (Number 2 grade fuel oil or 
Number 2 diesel fuel) are acceptable if the nominal fuel flow rate, 
temperature, and heat flux measurements conform to the requirements of 
paragraph VII(e) of this appendix.
    (vii) Fuel pressure regulator. Provide a fuel pressure regulator, 
adjusted to deliver a nominal 6.0 gal/hr (0.378 L/min) flow rate. An 
operating fuel pressure of 100 lb/in\2\ (0.71 MPa) for a nominally rated 
6.0 gal/hr 80[deg] spray angle nozzle (such as a PL type) delivers 6.0 
0.2 gal/hr (0.378 0.0126 L/
min).

[[Page 602]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.014

    (3) Calibration rig and equipment. (i) Construct individual 
calibration rigs to incorporate a calorimeter and thermocouple rake for 
the measurement of heat flux and temperature. Position the calibration 
rigs to allow movement of the burner from the test rig position to 
either the heat flux or temperature position with minimal difficulty.

[[Page 603]]

    (ii) Calorimeter. The calorimeter must be a total heat flux, foil 
type Gardon Gage of an appropriate range such as 0-20 Btu/ft \2\-sec (0-
22.7 W/cm \2\), accurate to 3% of the indicated 
reading. The heat flux calibration method must be in accordance with 
paragraph VI(b)(7) of this appendix.
    (iii) Calorimeter mounting. Mount the calorimeter in a 6- by 12- 
0.125 inch (152- by 305- 3 
mm) by 0.75 0.125 inch (19 mm 3 mm) thick insulating block which is attached to the 
heat flux calibration rig during calibration (figure 4). Monitor the 
insulating block for deterioration and replace it when necessary. Adjust 
the mounting as necessary to ensure that the calorimeter face is 
parallel to the exit plane of the test burner cone.

[[Page 604]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.015


[[Page 605]]


[GRAPHIC] [TIFF OMITTED] TR31JY03.016

    (iv) Thermocouples. Provide seven \1/8\-inch (3.2 mm) ceramic 
packed, metal sheathed, type K (Chromel-alumel), grounded junction 
thermocouples with a nominal 24 American Wire Gauge (AWG) size conductor 
for calibration. Attach the thermocouples to a steel angle bracket to 
form a thermocouple rake

[[Page 606]]

for placement in the calibration rig during burner calibration (figure 
5).
    (v) Air velocity meter. Use a vane-type air velocity meter to 
calibrate the velocity of air entering the burner. An Omega Engineering 
Model HH30A is satisfactory. Use a suitable adapter to attach the 
measuring device to the inlet side of the burner to prevent air from 
entering the burner other than through the measuring device, which would 
produce erroneously low readings. Use a flexible duct, measuring 4 
inches wide (102 mm) by 20 feet long (6.1 meters), to supply fresh air 
to the burner intake to prevent damage to the air velocity meter from 
ingested soot. An optional airbox permanently mounted to the burner 
intake area can effectively house the air velocity meter and provide a 
mounting port for the flexible intake duct.
    (4) Test specimen mounting frame. Make the mounting frame for the 
test specimens of \1/8\-inch (3.2 mm) thick steel as shown in figure 1, 
except for the center vertical former, which should be \1/4\-inch (6.4 
mm) thick to minimize warpage. The specimen mounting frame stringers 
(horizontal) should be bolted to the test frame formers (vertical) such 
that the expansion of the stringers will not cause the entire structure 
to warp. Use the mounting frame for mounting the two insulation blanket 
test specimens as shown in figure 2.
    (5) Backface calorimeters. Mount two total heat flux Gardon type 
calorimeters behind the insulation test specimens on the back side 
(cold) area of the test specimen mounting frame as shown in figure 6. 
Position the calorimeters along the same plane as the burner cone 
centerline, at a distance of 4 inches (102 mm) from the vertical 
centerline of the test frame.

[[Page 607]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.017

    (i) The calorimeters must be a total heat flux, foil type Gardon 
Gage of an appropriate range such as 0-5 Btu/ft\2\-sec (0-5.7 W/cm\2\), 
accurate to 3% of the indicated reading. The heat 
flux calibration method must comply with paragraph VI(b)(7) of this 
appendix.

[[Page 608]]

    (6) Instrumentation. Provide a recording potentiometer or other 
suitable calibrated instrument with an appropriate range to measure and 
record the outputs of the calorimeter and the thermocouples.
    (7) Timing device. Provide a stopwatch or other device, accurate to 
1%, to measure the time of application of the 
burner flame and burnthrough time.
    (8) Test chamber. Perform tests in a suitable chamber to reduce or 
eliminate the possibility of test fluctuation due to air movement. The 
chamber must have a minimum floor area of 10 by 10 feet (305 by 305 cm).
    (i) Ventilation hood. Provide the test chamber with an exhaust 
system capable of removing the products of combustion expelled during 
tests.
    (c) Test Specimens. (1) Specimen preparation. Prepare a minimum of 
three specimen sets of the same construction and configuration for 
testing.
    (2) Insulation blanket test specimen.
    (i) For batt-type materials such as fiberglass, the constructed, 
finished blanket specimen assemblies must be 32 inches wide by 36 inches 
long (81.3 by 91.4 cm), exclusive of heat sealed film edges.
    (ii) For rigid and other non-conforming types of insulation 
materials, the finished test specimens must fit into the test rig in 
such a manner as to replicate the actual in-service installation.
    (3) Construction. Make each of the specimens tested using the 
principal components (i.e., insulation, fire barrier material if used, 
and moisture barrier film) and assembly processes (representative seams 
and closures).
    (i) Fire barrier material. If the insulation blanket is constructed 
with a fire barrier material, place the fire barrier material in a 
manner reflective of the installed arrangement For example, if the 
material will be placed on the outboard side of the insulation material, 
inside the moisture film, place it the same way in the test specimen.
    (ii) Insulation material. Blankets that utilize more than one 
variety of insulation (composition, density, etc.) must have specimen 
sets constructed that reflect the insulation combination used. If, 
however, several blanket types use similar insulation combinations, it 
is not necessary to test each combination if it is possible to bracket 
the various combinations.
    (iii) Moisture barrier film. If a production blanket construction 
utilizes more than one type of moisture barrier film, perform separate 
tests on each combination. For example, if a polyimide film is used in 
conjunction with an insulation in order to enhance the burnthrough 
capabilities, also test the same insulation when used with a polyvinyl 
fluoride film.
    (iv) Installation on test frame. Attach the blanket test specimens 
to the test frame using 12 steel spring type clamps as shown in figure 
7. Use the clamps to hold the blankets in place in both of the outer 
vertical formers, as well as the center vertical former (4 clamps per 
former). The clamp surfaces should measure 1 inch by 2 inches (25 by 51 
mm). Place the top and bottom clamps 6 inches (15.2 cm) from the top and 
bottom of the test frame, respectively. Place the middle clamps 8 inches 
(20.3 cm) from the top and bottom clamps.

[[Page 609]]

[GRAPHIC] [TIFF OMITTED] TR31JY03.018

    (Note: For blanket materials that cannot be installed in accordance 
with figure 7 above, the blankets must be installed in a manner approved 
by the FAA.)
    (v) Conditioning. Condition the specimens at 70[deg] 5 [deg]F (21[deg] 2 [deg]C) and 
55% 10% relative humidity for a minimum of 24 
hours prior to testing.
    (d) Preparation of apparatus. (1) Level and center the frame 
assembly to ensure alignment of the calorimeter and/or thermocouple rake 
with the burner cone.
    (2) Turn on the ventilation hood for the test chamber. Do not turn 
on the burner blower. Measure the airflow of the test chamber using a 
vane anemometer or equivalent measuring device. The vertical air 
velocity just behind the top of the upper insulation blanket test 
specimen must be 100 50 ft/min (0.51 0.25 m/s). The horizontal air velocity at this point 
must be less than 50 ft/min (0.25 m/s).
    (3) If a calibrated flow meter is not available, measure the fuel 
flow rate using a graduated cylinder of appropriate size. Turn on the 
burner motor/fuel pump, after insuring that the igniter system is turned 
off. Collect the fuel via a plastic or rubber tube into the

[[Page 610]]

graduated cylinder for a 2-minute period. Determine the flow rate in 
gallons per hour. The fuel flow rate must be 6.0 0.2 gallons per hour (0.378 0.0126 
L/min).
    (e) Calibration. (1) Position the burner in front of the calorimeter 
so that it is centered and the vertical plane of the burner cone exit is 
4 0.125 inches (102 3 mm) 
from the calorimeter face. Ensure that the horizontal centerline of the 
burner cone is offset 1 inch below the horizontal centerline of the 
calorimeter (figure 8). Without disturbing the calorimeter position, 
rotate the burner in front of the thermocouple rake, such that the 
middle thermocouple (number 4 of 7) is centered on the burner cone.
[GRAPHIC] [TIFF OMITTED] TR31JY03.019


[[Page 611]]


    Ensure that the horizontal centerline of the burner cone is also 
offset 1 inch below the horizontal centerline of the thermocouple tips. 
Re-check measurements by rotating the burner to each position to ensure 
proper alignment between the cone and the calorimeter and thermocouple 
rake. (Note: The test burner mounting system must incorporate 
``detents'' that ensure proper centering of the burner cone with respect 
to both the calorimeter and the thermocouple rakes, so that rapid 
positioning of the burner can be achieved during the calibration 
procedure.)
    (2) Position the air velocity meter in the adapter or airbox, making 
certain that no gaps exist where air could leak around the air velocity 
measuring device. Turn on the blower/motor while ensuring that the fuel 
solenoid and igniters are off. Adjust the air intake velocity to a level 
of 2150 ft/min, (10.92 m/s) then turn off the blower/motor. (Note: The 
Omega HH30 air velocity meter measures 2.625 inches in diameter. To 
calculate the intake airflow, multiply the cross-sectional area (0.03758 
ft\2\) by the air velocity (2150 ft/min) to obtain 80.80 ft\3\/min. An 
air velocity meter other than the HH30 unit can be used, provided the 
calculated airflow of 80.80 ft\3\/min (2.29 m\3\/min) is equivalent.)
    (3) Rotate the burner from the test position to the warm-up 
position. Prior to lighting the burner, ensure that the calorimeter face 
is clean of soot deposits, and there is water running through the 
calorimeter. Examine and clean the burner cone of any evidence of 
buildup of products of combustion, soot, etc. Soot buildup inside the 
burner cone may affect the flame characteristics and cause calibration 
difficulties. Since the burner cone may distort with time, dimensions 
should be checked periodically.
    (4) While the burner is still rotated to the warm-up position, turn 
on the blower/motor, igniters and fuel flow, and light the burner. Allow 
it to warm up for a period of 2 minutes. Move the burner into the 
calibration position and allow 1 minute for calorimeter stabilization, 
then record the heat flux once every second for a period of 30 seconds. 
Turn off burner, rotate out of position, and allow to cool. Calculate 
the average heat flux over this 30-second duration. The average heat 
flux should be 16.0 0.8 Btu/ft\2\ sec (18.2 0.9 W/cm\2\).
    (5) Position the burner in front of the thermocouple rake. After 
checking for proper alignment, rotate the burner to the warm-up 
position, turn on the blower/motor, igniters and fuel flow, and light 
the burner. Allow it to warm up for a period of 2 minutes. Move the 
burner into the calibration position and allow 1 minute for thermocouple 
stabilization, then record the temperature of each of the 7 
thermocouples once every second for a period of 30 seconds. Turn off 
burner, rotate out of position, and allow to cool. Calculate the average 
temperature of each thermocouple over this 30-second period and record. 
The average temperature of each of the 7 thermocouples should be 1900 
[deg]F 100 [deg]F (1038 56 
[deg]C).
    (6) If either the heat flux or the temperatures are not within the 
specified range, adjust the burner intake air velocity and repeat the 
procedures of paragraphs (4) and (5) above to obtain the proper values. 
Ensure that the inlet air velocity is within the range of 2150 ft/min 
50 ft/min (10.92 0.25 m/s).
    (7) Calibrate prior to each test until consistency has been 
demonstrated. After consistency has been confirmed, several tests may be 
conducted with calibration conducted before and after a series of tests.
    (f) Test procedure. (1) Secure the two insulation blanket test 
specimens to the test frame. The insulation blankets should be attached 
to the test rig center vertical former using four spring clamps 
positioned as shown in figure 7 (according to the criteria of paragraph 
paragraph (c)(3)(iv) of this part of this appendix).
    (2) Ensure that the vertical plane of the burner cone is at a 
distance of 4 0.125 inch (102 3 mm) from the outer surface of the horizontal stringers 
of the test specimen frame, and that the burner and test frame are both 
situated at a 30[deg] angle with respect to vertical.
    (3) When ready to begin the test, direct the burner away from the 
test position to the warm-up position so that the flame will not impinge 
on the specimens prematurely. Turn on and light the burner and allow it 
to stabilize for 2 minutes.
    (4) To begin the test, rotate the burner into the test position and 
simultaneously start the timing device.
    (5) Expose the test specimens to the burner flame for 4 minutes and 
then turn off the burner. Immediately rotate the burner out of the test 
position.
    (6) Determine (where applicable) the burnthrough time, or the point 
at which the heat flux exceeds 2.0 Btu/ft\2\-sec (2.27 W/cm\2\).
    (g) Report. (1) Identify and describe the specimen being tested.
    (2) Report the number of insulation blanket specimens tested.
    (3) Report the burnthrough time (if any), and the maximum heat flux 
on the back face of the insulation blanket test specimen, and the time 
at which the maximum occurred.
    (h) Requirements. (1) Each of the two insulation blanket test 
specimens must not allow fire or flame penetration in less than 4 
minutes.
    (2) Each of the two insulation blanket test specimens must not allow 
more than 2.0 Btu/

[[Page 612]]

ft\2\-sec (2.27 W/cm\2\) on the cold side of the insulation specimens at 
a point 12 inches (30.5 cm) from the face of the test rig.

[Amdt. 25-32, 37 FR 3972, Feb. 24, 1972]

    Editorial Note: For Federal Register citations affecting appendix F 
to Part 25, see the List of CFR Sections Affected, which appears in the 
Finding Aids section of the printed volume and at www.fdsys.gov.



       Sec. Appendix G to Part 25--Continuous Gust Design Criteria

    The continuous gust design criteria in this appendix must be used in 
establishing the dynamic response of the airplane to vertical and 
lateral continuous turbulence unless a more rational criteria is used. 
The following gust load requirements apply to mission analysis and 
design envelope analysis:
    (a) The limit gust loads utilizing the continuous turbulence concept 
must be determined in accordance with the provisions of either paragraph 
(b) or paragraphs (c) and (d) of this appendix.
    (b) Design envelope analysis. The limit loads must be determined in 
accordance with the following:
    (1) All critical altitudes, weights, and weight distributions, as 
specified in Sec. 25.321(b), and all critical speeds within the ranges 
indicated in paragraph (b)(3) of this appendix must be considered.
    (2) Values of [Amacr] (ratio of root-mean-square incremental load 
root-mean-square gust velocity) must be determined by dynamic analysis. 
The power spectral density of the atmospheric turbulence must be as 
given by the equation--
[GRAPHIC] [TIFF OMITTED] TC28SE91.078

where:

[phis]=power-spectral density (ft./sec.) \2\/rad./ft.
[sigma]=root-mean-square gust velocity, ft./sec.
[Omega]=reduced frequency, radians per foot.
L=2,500 ft.

    (3) The limit loads must be obtained by multiplying the [Amacr] 
values determined by the dynamic analysis by the following values of the 
gust velocity U[sigma]:
    (i) At speed Vc: U[sigma]=85 fps true gust velocity in 
the interval 0 to 30,000 ft. altitude and is linearly decreased to 30 
fps true gust velocity at 80,000 ft. altitude. Where the Administrator 
finds that a design is comparable to a similar design with extensive 
satisfactory service experience, it will be acceptable to select 
U[sigma] at Vc less than 85 fps, but not less than 75 fps, 
with linear decrease from that value at 20,000 feet to 30 fps at 80,000 
feet. The following factors will be taken into account when assessing 
comparability to a similar design:
    (1) The transfer function of the new design should exhibit no 
unusual characteristics as compared to the similar design which will 
significantly affect response to turbulence; e.g., coalescence of modal 
response in the frequency regime which can result in a significant 
increase of loads.
    (2) The typical mission of the new airplane is substantially 
equivalent to that of the similar design.
    (3) The similar design should demonstrate the adequacy of the 
U[sigma] selected.
    (ii) At speed VB: U[sigma] is equal to 1.32 times the 
values obtained under paragraph (b)(3)(i) of this appendix.
    (iii) At speed VD: U[sigma] is equal to \1/2\ the values 
obtained under paragraph (b)(3)(i) of this appendix.
    (iv) At speeds between VB and Vc and between 
Vc and VD: U[sigma] is equal to a value obtained 
by linear interpolation.
    (4) When a stability augmentation system is included in the 
analysis, the effect of system nonlinearities on loads at the limit load 
level must be realistically or conservatively accounted for.
    (c) Mission analysis. Limit loads must be determined in accordance 
with the following:
    (1) The expected utilization of the airplane must be represented by 
one or more flight profiles in which the load distribution and the 
variation with time of speed, altitude, gross weight, and center of 
gravity position are defined. These profiles must be divided into 
mission segments or blocks, for analysis, and average or effective 
values of the pertinent parameters defined for each segment.
    (2) For each of the mission segments defined under paragraph (c)(1) 
of this appendix, values of [Amacr] and No must be determined 
by analysis. [Amacr] is defined as the ratio of root-mean-square 
incremental load to root-mean-square gust velocity and No is 
the radius of gyration of the load power spectral density function about 
zero frequency. The power spectral density of the atmospheric turbulence 
must be given by the equation set forth in paragraph (b)(2) of this 
appendix.
    (3) For each of the load and stress quantities selected, the 
frequency of exceedance must be determined as a function of load level 
by means of the equation--

[[Page 613]]

[GRAPHIC] [TIFF OMITTED] TC28SE91.079

where--

t=selected time interval.
y=net value of the load or stress.
Yone=g=value of the load or stress in one-g level flight.
N(y)=average number of exceedances of the indicated value of the load or 
          stress in unit time.
[sum]=symbol denoting summation over all mission segments.
No, [Amacr]=parameters determined by dynamic analysis as 
          defined in paragraph (c)(2) of this appendix.
P1, P2, b1, b2=parameters 
          defining the probability distributions of root-mean-square 
          gust velocity, to be read from Figures 1 and 2 of this 
          appendix.

The limit gust loads must be read from the frequency of exceedance 
curves at a frequency of exceedance of 2x10-5 exceedances per hour. Both 
positive and negative load directions must be considered in determining 
the limit loads.
    (4) If a stability augmentation system is utilized to reduce the 
gust loads, consideration must be given to the fraction of flight time 
that the system may be inoperative. The flight profiles of paragraph 
(c)(1) of this appendix must include flight with the system inoperative 
for this fraction of the flight time. When a stability augmentation 
system is included in the analysis, the effect of system nonlinearities 
on loads at the limit load level must be conservatively accounted for.
    (d) Supplementary design envelope analysis. In addition to the limit 
loads defined by paragraph (c) of this appendix, limit loads must also 
be determined in accordance with paragraph (b) of this appendix, except 
that--
    (1) In paragraph (b)(3)(i) of this appendix, the value of 
U[sigma]=85 fps true gust velocity is replaced by U[sigma]=60 fps true 
gust velocity on the interval 0 to 30,000 ft. altitude, and is linearly 
decreased to 25 fps true gust velocity at 80,000 ft. altitude; and
    (2) In paragraph (b) of this appendix, the reference to paragraphs 
(b)(3)(i) through (b)(3)(iii) of this appendix is to be understood as 
referring to the paragraph as modified by paragraph (d)(1).

[[Page 614]]

[GRAPHIC] [TIFF OMITTED] TC28SE91.080


[[Page 615]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.081


[Amdt. 25-54, 45 FR 60173, Sept. 11, 1980]

    Effective Date Note: At 79 FR 73469, Dec. 11, 2014, Appendix G to 
part 25 was removed and reserved, effective Feb. 9, 2015.



  Sec. Appendix H to Part 25--Instructions for Continued Airworthiness

H25.1 General.

    (a) This appendix specifies requirements for preparation of 
Instructions for Continued Airworthiness as required by Sec. Sec. 
25.1529, 25.1729, and applicable provisions of parts 21 and 26 of this 
chapter.
    (b) The Instructions for Continued Airworthiness for each airplane 
must include the Instructions for Continued Airworthiness for each 
engine and propeller (hereinafter designated ``products''), for each 
appliance required by this chapter, and any required information 
relating to the interface of

[[Page 616]]

those appliances and products with the airplane. If Instructions for 
Continued Airworthiness are not supplied by the manufacturer of an 
appliance or product installed in the airplane, the Instructions for 
Continued Airworthiness for the airplane must include the information 
essential to the continued airworthiness of the airplane.
    (c) The applicant must submit to the FAA a program to show how 
changes to the Instructions for Continued Airworthiness made by the 
applicant or by the manufacturers or products and appliances installed 
in the airplane will be distributed.

H25.2 Format.

    (a) The Instructions for Continued Airworthiness must be in the form 
of a manual or manuals as appropriate for the quantity of data to be 
provided.
    (b) The format of the manual or manuals must provide for a practical 
arrangement.

H25.3 Content.
    The contents of the manual or manuals must be prepared in the 
English language. The Instructions for Continued Airworthiness must 
contain the following manuals or sections, as appropriate, and 
information:
    (a) Airplane maintenance manual or section. (1) Introduction 
information that includes an explanation of the airplane's features and 
data to the extent necessary for maintenance or preventive maintenance.
    (2) A description of the airplane and its systems and installations 
including its engines, propellers, and appliances.
    (3) Basic control and operation information describing how the 
airplane components and systems are controlled and how they operate, 
including any special procedures and limitations that apply.
    (4) Servicing information that covers details regarding servicing 
points, capacities of tanks, reservoirs, types of fluids to be used, 
pressures applicable to the various systems, location of access panels 
for inspection and servicing, locations of lubrication points, 
lubricants to be used, equipment required for servicing, tow 
instructions and limitations, mooring, jacking, and leveling 
information.
    (b) Maintenance instructions. (1) Scheduling information for each 
part of the airplane and its engines, auxiliary power units, propellers, 
accessories, instruments, and equipment that provides the recommended 
periods at which they should be cleaned, inspected, adjusted, tested, 
and lubricated, and the degree of inspection, the applicable wear 
tolerances, and work recommended at these periods. However, the 
applicant may refer to an accessory, instrument, or equipment 
manufacturer as the source of this information if the applicant shows 
that the item has an exceptionally high degree of complexity requiring 
specialized maintenance techniques, test equipment, or expertise. The 
recommended overhaul periods and necessary cross references to the 
Airworthiness Limitations section of the manual must also be included. 
In addition, the applicant must include an inspection program that 
includes the frequency and extent of the inspections necessary to 
provide for the continued airworthiness of the airplane.
    (2) Troubleshooting information describing probable malfunctions, 
how to recognize those malfunctions, and the remedial action for those 
malfunctions.
    (3) Information describing the order and method of removing and 
replacing products and parts with any necessary precautions to be taken.
    (4) Other general procedural instructions including procedures for 
system testing during ground running, symmetry checks, weighing and 
determining the center of gravity, lifting and shoring, and storage 
limitations.
    (c) Diagrams of structural access plates and information needed to 
gain access for inspections when access plates are not provided.
    (d) Details for the application of special inspection techniques 
including radiographic and ultrasonic testing where such processes are 
specified.
    (e) Information needed to apply protective treatments to the 
structure after inspection.
    (f) All data relative to structural fasteners such as 
identification, discard recommendations, and torque values.
    (g) A list of special tools needed.

H25.4 Airworthiness Limitations section.

    (a) The Instructions for Continued Airworthiness must contain a 
section titled Airworthiness Limitations that is segregated and clearly 
distinguishable from the rest of the document. This section must set 
forth--
    (1) Each mandatory modification time, replacement time, structural 
inspection interval, and related structural inspection procedure 
approved under Sec. 25.571.
    (2) Each mandatory replacement time, inspection interval, related 
inspection procedure, and all critical design configuration control 
limitations approved under Sec. 25.981 for the fuel tank system.
    (3) Any mandatory replacement time of EWIS components as defined in 
section 25.1701.
    (4) A limit of validity of the engineering data that supports the 
structural maintenance program (LOV), stated as a total number of 
accumulated flight cycles or flight hours or both, approved under Sec. 
25.571. Until the full-scale fatigue testing is completed and the FAA 
has approved the LOV, the number of cycles accumulated by the airplane 
cannot be greater than \1/2\ the number of cycles accumulated on the 
fatigue test article.
    (b) If the Instructions for Continued Airworthiness consist of 
multiple documents,

[[Page 617]]

the section required by this paragraph must be included in the principal 
manual. This section must contain a legible statement in a prominent 
location that reads: ``The Airworthiness Limitations section is FAA-
approved and specifies maintenance required under Sec. Sec. 43.16 and 
91.403 of the Federal Aviation Regulations, unless an alternative 
program has been FAA approved.''
    H25.5 Electrical Wiring Interconnection System (EWIS) Instructions 
for Continued Airworthiness.
    (a) The applicant must prepare Instructions for Continued 
Airworthiness (ICA) applicable to EWIS as defined by Sec. 25.1701 that 
are approved by the FAA and include the following:
    (1) Maintenance and inspection requirements for the EWIS developed 
with the use of an enhanced zonal analysis procedure that includes:
    (i) Identification of each zone of the airplane.
    (ii) Identification of each zone that contains EWIS.
    (iii) Identification of each zone containing EWIS that also contains 
combustible materials.
    (iv) Identification of each zone in which EWIS is in close proximity 
to both primary and back-up hydraulic, mechanical, or electrical flight 
controls and lines.
    (v) Identification of--
    (A) Tasks, and the intervals for performing those tasks, that will 
reduce the likelihood of ignition sources and accumulation of 
combustible material, and
    (B) Procedures, and the intervals for performing those procedures, 
that will effectively clean the EWIS components of combustible material 
if there is not an effective task to reduce the likelihood of 
combustible material accumulation.
    (vi) Instructions for protections and caution information that will 
minimize contamination and accidental damage to EWIS, as applicable, 
during performance of maintenance, alteration, or repairs.
    (2) Acceptable EWIS maintenance practices in a standard format.
    (3) Wire separation requirements as determined under Sec. 25.1707.
    (4) Information explaining the EWIS identification method and 
requirements for identifying any changes to EWIS under Sec. 25.1711.
    (5) Electrical load data and instructions for updating that data.
    (b) The EWIS ICA developed in accordance with the requirements of 
H25.5(a)(1) must be in the form of a document appropriate for the 
information to be provided, and they must be easily recognizable as EWIS 
ICA. This document must either contain the required EWIS ICA or 
specifically reference other portions of the ICA that contain this 
information.

[Amdt. 25-54, 45 FR 60177, Sept. 11, 1980, as amended by Amdt. 25-68, 54 
FR 34329, Aug. 18, 1989; Amdt. 25-102, 66 FR 23130, May 7, 2001; Amdt. 
25-123, 72 FR 63408, Nov. 8, 2007; Amdt. 25-132, 75 FR 69782, Nov. 15, 
2010]



Sec. Appendix I to Part 25--Installation of an Automatic Takeoff Thrust 
                         Control System (ATTCS)

I25.1 General.

    (a) This appendix specifies additional requirements for installation 
of an engine power control system that automatically resets thrust or 
power on operating engine(s) in the event of any one engine failure 
during takeoff.
    (b) With the ATTCS and associated systems functioning normally as 
designed, all applicable requirements of Part 25, except as provided in 
this appendix, must be met without requiring any action by the crew to 
increase thrust or power.

I25.2 Definitions.

    (a) Automatic Takeoff Thrust Control System (ATTCS). An ATTCS is 
defined as the entire automatic system used on takeoff, including all 
devices, both mechanical and electrical, that sense engine failure, 
transmit signals, actuate fuel controls or power levers or increase 
engine power by other means on operating engines to achieve scheduled 
thrust or power increases, and furnish cockpit information on system 
operation.
    (b) Critical Time Interval. When conducting an ATTCS takeoff, the 
critical time interval is between V1 minus 1 second and a 
point on the minimum performance, all-engine flight path where, assuming 
a simultaneous occurrence of an engine and ATTCS failure, the resulting 
minimum flight path thereafter intersects the Part 25 required actual 
flight path at no less than 400 feet above the takeoff surface. This 
time interval is shown in the following illustration:

[[Page 618]]

[GRAPHIC] [TIFF OMITTED] TC28SE91.082

I25.3 Performance and System Reliability Requirements.

    The applicant must comply with the performance and ATTCS reliability 
requirements as follows:
    (a) An ATTCS failure or a combination of failures in the ATTCS 
during the critical time interval:
    (1) Shall not prevent the insertion of the maximum approved takeoff 
thrust or power, or must be shown to be an improbable event.
    (2) Shall not result in a significant loss or reduction in thrust or 
power, or must be shown to be an extremely improbable event.
    (b) The concurrent existence of an ATTCS failure and an engine 
failure during the critical time interval must be shown to be extremely 
improbable.
    (c) All applicable performance requirements of Part 25 must be met 
with an engine failure occurring at the most critical point during 
takeoff with the ATTCS system functioning.

I25.4 Thrust Setting.

    The initial takeoff thrust or power setting on each engine at the 
beginning of the takeoff roll may not be less than any of the following:
    (a) Ninety (90) percent of the thrust or power set by the ATTCS (the 
maximum takeoff thrust or power approved for the airplane under existing 
ambient conditions);
    (b) That required to permit normal operation of all safety-related 
systems and equipment dependent upon engine thrust or power lever 
position; or
    (c) That shown to be free of hazardous engine response 
characteristics when thrust or power is advanced from the initial 
takeoff thrust or power to the maximum approved takeoff thrust or power.

I25.5 Powerplant Controls.


[[Page 619]]


    (a) In addition to the requirements of Sec. 25.1141, no single 
failure or malfunction, or probable combination thereof, of the ATTCS, 
including associated systems, may cause the failure of any powerplant 
function necessary for safety.
    (b) The ATTCS must be designed to:
    (1) Apply thrust or power on the operating engine(s), following any 
one engine failure during takeoff, to achieve the maximum approved 
takeoff thrust or power without exceeding engine operating limits;
    (2) Permit manual decrease or increase in thrust or power up to the 
maximum takeoff thrust or power approved for the airplane under existing 
conditions through the use of the power lever. For airplanes equipped 
with limiters that automatically prevent engine operating limits from 
being exceeded under existing ambient conditions, other means may be 
used to increase the thrust or power in the event of an ATTCS failure 
provided the means is located on or forward of the power levers; is 
easily identified and operated under all operating conditions by a 
single action of either pilot with the hand that is normally used to 
actuate the power levers; and meets the requirements of Sec. 25.777 
(a), (b), and (c);
    (3) Provide a means to verify to the flightcrew before takeoff that 
the ATTCS is in a condition to operate; and
    (4) Provide a means for the flightcrew to deactivate the automatic 
function. This means must be designed to prevent inadvertent 
deactivation.

I25.6 Powerplant Instruments.

    In addition to the requirements of Sec. 25.1305:
    (a) A means must be provided to indicate when the ATTCS is in the 
armed or ready condition; and
    (b) If the inherent flight characteristics of the airplane do not 
provide adequate warning that an engine has failed, a warning system 
that is independent of the ATTCS must be provided to give the pilot a 
clear warning of any engine failure during takeoff.

[Amdt. 25-62, 52 FR 43156, Nov. 9, 1987]



            Sec. Appendix J to Part 25--Emergency Evacuation

    The following test criteria and procedures must be used for showing 
compliance with Sec. 25.803:
    (a) The emergency evacuation must be conducted with exterior ambient 
light levels of no greater than 0.3 foot-candles prior to the activation 
of the airplane emergency lighting system. The source(s) of the initial 
exterior ambient light level may remain active or illuminated during the 
actual demonstration. There must, however, be no increase in the 
exterior ambient light level except for that due to activation of the 
airplane emergency lighting system.
    (b) The airplane must be in a normal attitude with landing gear 
extended.
    (c) Unless the airplane is equipped with an off-wing descent means, 
stands or ramps may be used for descent from the wing to the ground. 
Safety equipment such as mats or inverted life rafts may be placed on 
the floor or ground to protect participants. No other equipment that is 
not part of the emergency evacuation equipment of the airplane may be 
used to aid the participants in reaching the ground.
    (d) Except as provided in paragraph (a) of this appendix, only the 
airplane's emergency lighting system may provide illumination.
    (e) All emergency equipment required for the planned operation of 
the airplane must be installed.
    (f) Each internal door or curtain must be in the takeoff 
configuration.
    (g) Each crewmember must be seated in the normally assigned seat for 
takeoff and must remain in the seat until receiving the signal for 
commencement of the demonstration. Each crewmember must be a person 
having knowledge of the operation of exits and emergency equipment and, 
if compliance with Sec. 121.291 is also being demonstrated, each flight 
attendant must be a member of a regularly scheduled line crew.
    (h) A representative passenger load of persons in normal health must 
be used as follows:
    (1) At least 40 percent of the passenger load must be female.
    (2) At least 35 percent of the passenger load must be over 50 years 
of age.
    (3) At least 15 percent of the passenger load must be female and 
over 50 years of age.
    (4) Three life-size dolls, not included as part of the total 
passenger load, must be carried by passengers to simulate live infants 2 
years old or younger.
    (5) Crewmembers, mechanics, and training personnel, who maintain or 
operate the airplane in the normal course of their duties, may not be 
used as passengers.
    (i) No passenger may be assigned a specific seat except as the 
Administrator may require. Except as required by subparagraph (g) of 
this paragraph, no employee of the applicant may be seated next to an 
emergency exit.
    (j) Seat belts and shoulder harnesses (as required) must be 
fastened.
    (k) Before the start of the demonstration, approximately one-half of 
the total average amount of carry-on baggage, blankets, pillows, and 
other similar articles must be distributed at several locations in 
aisles and emergency exit access ways to create minor obstructions.
    (l) No prior indication may be given to any crewmember or passenger 
of the particular exits to be used in the demonstration.
    (m) The applicant may not practice, rehearse, or describe the 
demonstration for the

[[Page 620]]

participants nor may any participant have taken part in this type of 
demonstration within the preceding 6 months.
    (n) Prior to entering the demonstration aircraft, the passengers may 
also be advised to follow directions of crewmembers but may not be 
instructed on the procedures to be followed in the demonstration, except 
with respect to safety procedures in place for the demonstration or 
which have to do with the demonstration site. Prior to the start of the 
demonstration, the pre-takeoff passenger briefing required by Sec. 
121.571 may be given. Flight attendants may assign demonstration 
subjects to assist persons from the bottom of a slide, consistent with 
their approved training program.
    (o) The airplane must be configured to prevent disclosure of the 
active emergency exits to demonstration participants in the airplane 
until the start of the demonstration.
    (p) Exits used in the demonstration must consist of one exit from 
each exit pair. The demonstration may be conducted with the escape 
slides, if provided, inflated and the exits open at the beginning of the 
demonstration. In this case, all exits must be configured such that the 
active exits are not disclosed to the occupants. If this method is used, 
the exit preparation time for each exit utilized must be accounted for, 
and exits that are not to be used in the demonstration must not be 
indicated before the demonstration has started. The exits to be used 
must be representative of all of the emergency exits on the airplane and 
must be designated by the applicant, subject to approval by the 
Administrator. At least one floor level exit must be used.
    (q) Except as provided in paragraph (c) of this section, all 
evacuees must leave the airplane by a means provided as part of the 
airplane's equipment.
    (r) The applicant's approved procedures must be fully utilized, 
except the flightcrew must take no active role in assisting others 
inside the cabin during the demonstration.
    (s) The evacuation time period is completed when the last occupant 
has evacuated the airplane and is on the ground. Provided that the 
acceptance rate of the stand or ramp is no greater than the acceptance 
rate of the means available on the airplane for descent from the wing 
during an actual crash situation, evacuees using stands or ramps allowed 
by paragraph (c) of this appendix are considered to be on the ground 
when they are on the stand or ramp.

[Amdt. 25-72, 55 FR 29788, July 20, 1990, as amended by Amdt. 25-79, 
Aug. 26, 1993; Amdt. 25-117, 69 FR 67499, Nov. 17, 2004]



         Sec. Appendix K to Part 25--Extended Operations (ETOPS)

    This appendix specifies airworthiness requirements for the approval 
of an airplane-engine combination for extended operations (ETOPS). For 
two-engine airplanes, the applicant must comply with sections K25.1 and 
K25.2 of this appendix. For airplanes with more than two engines, the 
applicant must comply with sections K25.1 and K25.3 of this appendix.
    K25.1 Design requirements.
    K25.1.1 Part 25 compliance.
    The airplane-engine combination must comply with the requirements of 
part 25 considering the maximum flight time and the longest diversion 
time for which the applicant seeks approval.
    K25.1.2 Human factors.
    An applicant must consider crew workload, operational implications, 
and the crew's and passengers' physiological needs during continued 
operation with failure effects for the longest diversion time for which 
it seeks approval.
    K25.1.3 Airplane systems.
    (a) Operation in icing conditions.
    (1) The airplane must be certificated for operation in icing 
conditions in accordance with Sec. 25.1419.
    (2) The airplane must be able to safely conduct an ETOPS diversion 
with the most critical ice accretion resulting from:
    (i) Icing conditions encountered at an altitude that the airplane 
would have to fly following an engine failure or cabin decompression.
    (ii) A 15-minute hold in the continuous maximum icing conditions 
specified in Appendix C of this part with a liquid water content factor 
of 1.0.
    (iii) Ice accumulated during approach and landing in the icing 
conditions specified in Appendix C of this part.
    (b) Electrical power supply. The airplane must be equipped with at 
least three independent sources of electrical power.
    (c) Time limited systems. The applicant must define the system time 
capability of each ETOPS significant system that is time-limited.
    K25.1.4 Propulsion systems.
    (a) Fuel system design. Fuel necessary to complete an ETOPS flight 
(including a diversion for the longest time for which the applicant 
seeks approval) must be available to the operating engines at the 
pressure and fuel-flow required by Sec. 25.955 under any airplane 
failure condition not shown to be extremely improbable. Types of 
failures that must be considered include, but are not limited to: 
crossfeed valve failures, automatic fuel management system failures, and 
normal electrical power generation failures.
    (1) If the engine has been certified for limited operation with 
negative engine-fuel-

[[Page 621]]

pump-inlet pressures, the following requirements apply:
    (i) Airplane demonstration-testing must cover worst case cruise and 
diversion conditions involving:
    (A) Fuel grade and temperature.
    (B) Thrust or power variations.
    (C) Turbulence and negative G.
    (D) Fuel system components degraded within their approved 
maintenance limits.
    (ii) Unusable-fuel quantity in the suction feed configuration must 
be determined in accordance with Sec. 25.959.
    (2) For two-engine airplanes to be certificated for ETOPS beyond 180 
minutes, one fuel boost pump in each main tank and at least one 
crossfeed valve, or other means for transferring fuel, must be powered 
by an independent electrical power source other than the three power 
sources required to comply with section K25.1.3(b) of this appendix. 
This requirement does not apply if the normal fuel boost pressure, 
crossfeed valve actuation, or fuel transfer capability is not provided 
by electrical power.
    (3) An alert must be displayed to the flightcrew when the quantity 
of fuel available to the engines falls below the level required to fly 
to the destination. The alert must be given when there is enough fuel 
remaining to safely complete a diversion. This alert must account for 
abnormal fuel management or transfer between tanks, and possible loss of 
fuel. This paragraph does not apply to airplanes with a required flight 
engineer.
    (b) APU design. If an APU is needed to comply with this appendix, 
the applicant must demonstrate that:
    (1) The reliability of the APU is adequate to meet those 
requirements; and
    (2) If it is necessary that the APU be able to start in flight, it 
is able to start at any altitude up to the maximum operating altitude of 
the airplane, or 45,000 feet, whichever is lower, and run for the 
remainder of any flight .
    (c) Engine oil tank design. The engine oil tank filler cap must 
comply with Sec. 33.71(c)(4) of this chapter.
    K25.1.5 Engine-condition monitoring.
    Procedures for engine-condition monitoring must be specified and 
validated in accordance with Part 33, Appendix A, paragraph A33.3(c) of 
this chapter.
    K25.1.6 Configuration, maintenance, and procedures.
    The applicant must list any configuration, operating and maintenance 
requirements, hardware life limits, MMEL constraints, and ETOPS approval 
in a CMP document.
    K25.1.7 Airplane flight manual.
    The airplane flight manual must contain the following information 
applicable to the ETOPS type design approval:
    (a) Special limitations, including any limitation associated with 
operation of the airplane up to the maximum diversion time being 
approved.
    (b) Required markings or placards.
    (c) The airborne equipment required for extended operations and 
flightcrew operating procedures for this equipment.
    (d) The system time capability for the following:
    (1) The most limiting fire suppression system for Class C cargo or 
baggage compartments.
    (2) The most limiting ETOPS significant system other than fire 
suppression systems for Class C cargo or baggage compartments.
    (e) This statement: ``The type-design reliability and performance of 
this airplane-engine combination has been evaluated under 14 CFR 25.1535 
and found suitable for (identify maximum approved diversion time) 
extended operations (ETOPS) when the configuration, maintenance, and 
procedures standard contained in (identify the CMP document) are met. 
The actual maximum approved diversion time for this airplane may be less 
based on its most limiting system time capability. This finding does not 
constitute operational approval to conduct ETOPS.''
    K25.2. Two-engine airplanes.
    An applicant for ETOPS type design approval of a two-engine airplane 
must use one of the methods described in section K25.2.1, K25.2.2, or 
K25.2.3 of this appendix.
    K25.2.1 Service experience method.
    An applicant for ETOPS type design approval using the service 
experience method must comply with sections K25.2.1(a) and K25.2.1(b) of 
this appendix before conducting the assessments specified in sections 
K25.2.1(c) and K25.2.1(d) of this appendix, and the flight test 
specified in section K25.2.1(e) of this appendix.
    (a) Service experience. The world fleet for the airplane-engine 
combination must accumulate a minimum of 250,000 engine-hours. The FAA 
may reduce this number of hours if the applicant identifies compensating 
factors that are acceptable to the FAA. The compensating factors may 
include experience on another airplane, but experience on the candidate 
airplane must make up a significant portion of the total service 
experience.
    (b) In-flight shutdown (IFSD) rates. The demonstrated 12-month 
rolling average IFSD rate for the world fleet of the airplane-engine 
combination must be commensurate with the level of ETOPS approval being 
sought.
    (1) For type design approval up to and including 120 minutes: An 
IFSD rate of 0.05 or less per 1,000 world-fleet engine-hours, unless 
otherwise approved by the FAA. Unless the IFSD rate is 0.02 or less per 
1,000 world-fleet engine-hours, the applicant must provide a

[[Page 622]]

list of corrective actions in the CMP document specified in section 
K25.1.6 of this appendix, that, when taken, would result in an IFSD rate 
of 0.02 or less per 1,000 fleet engine-hours.
    (2) For type design approval up to and including 180 minutes: An 
IFSD rate of 0.02 or less per 1,000 world-fleet engine-hours, unless 
otherwise approved by the FAA. If the airplane-engine combination does 
not meet this rate by compliance with an existing 120-minute CMP 
document, then new or additional CMP requirements that the applicant has 
demonstrated would achieve this IFSD rate must be added to the CMP 
document.
    (3) For type design approval beyond 180 minutes: An IFSD rate of 
0.01 or less per 1,000 fleet engine-hours unless otherwise approved by 
the FAA. If the airplane-engine combination does not meet this rate by 
compliance with an existing 120-minute or 180-minute CMP document, then 
new or additional CMP requirements that the applicant has demonstrated 
would achieve this IFSD rate must be added to the CMP document.
    (c) Propulsion system assessment. (1) The applicant must conduct a 
propulsion system assessment based on the following data collected from 
the world-fleet of the airplane-engine combination:
    (i) A list of all IFSD's, unplanned ground engine shutdowns, and 
occurrences (both ground and in-flight) when an engine was not shut 
down, but engine control or the desired thrust or power level was not 
achieved, including engine flameouts. Planned IFSD's performed during 
flight training need not be included. For each item, the applicant must 
provide--
    (A) Each airplane and engine make, model, and serial number;
    (B) Engine configuration, and major alteration history;
    (C) Engine position;
    (D) Circumstances leading up to the engine shutdown or occurrence;
    (E) Phase of flight or ground operation;
    (F) Weather and other environmental conditions; and
    (G) Cause of engine shutdown or occurrence.
    (ii) A history of unscheduled engine removal rates since 
introduction into service (using 6- and 12-month rolling averages), with 
a summary of the major causes for the removals.
    (iii) A list of all propulsion system events (whether or not caused 
by maintenance or flightcrew error), including dispatch delays, 
cancellations, aborted takeoffs, turnbacks, diversions, and flights that 
continue to destination after the event.
    (iv) The total number of engine hours and cycles, the number of 
hours for the engine with the highest number of hours, the number of 
cycles for the engine with the highest number of cycles, and the 
distribution of hours and cycles.
    (v) The mean time between failures (MTBF) of propulsion system 
components that affect reliability.
    (vi) A history of the IFSD rates since introduction into service 
using a 12-month rolling average.
    (2) The cause or potential cause of each item listed in 
K25.2.1(c)(1)(i) must have a corrective action or actions that are shown 
to be effective in preventing future occurrences. Each corrective action 
must be identified in the CMP document specified in section K25.1.6. A 
corrective action is not required:
    (i) For an item where the manufacturer is unable to determine a 
cause or potential cause.
    (ii) For an event where it is technically unfeasible to develop a 
corrective action.
    (iii) If the world-fleet IFSD rate--
    (A) Is at or below 0.02 per 1,000 world-fleet engine-hours for 
approval up to and including 180-minute ETOPS; or
    (B) Is at or below 0.01 per 1,000 world-fleet engine-hours for 
approval greater than 180-minute ETOPS.
    (d) Airplane systems assessment. The applicant must conduct an 
airplane systems assessment. The applicant must show that the airplane 
systems comply with Sec. 25.1309(b) using available in-service 
reliability data for ETOPS significant systems on the candidate 
airplane-engine combination. Each cause or potential cause of a relevant 
design, manufacturing, operational, and maintenance problem occurring in 
service must have a corrective action or actions that are shown to be 
effective in preventing future occurrences. Each corrective action must 
be identified in the CMP document specified in section K25.1.6 of this 
appendix. A corrective action is not required if the problem would not 
significantly impact the safety or reliability of the airplane system 
involved. A relevant problem is a problem with an ETOPS group 1 
significant system that has or could result in, an IFSD or diversion. 
The applicant must include in this assessment relevant problems with 
similar or identical equipment installed on other types of airplanes to 
the extent such information is reasonably available.
    (e) Airplane flight test. The applicant must conduct a flight test 
to validate the flightcrew's ability to safely conduct an ETOPS 
diversion with an inoperative engine and worst-case ETOPS Significant 
System failures and malfunctions that could occur in service. The flight 
test must validate the airplane's flying qualities and performance with 
the demonstrated failures and malfunctions.
    K25.2.2 Early ETOPS method.
    An applicant for ETOPS type design approval using the Early ETOPS 
method must comply with the following requirements:

[[Page 623]]

    (a) Assessment of relevant experience with airplanes previously 
certificated under part 25. The applicant must identify specific 
corrective actions taken on the candidate airplane to prevent relevant 
design, manufacturing, operational, and maintenance problems experienced 
on airplanes previously certificated under part 25 manufactured by the 
applicant. Specific corrective actions are not required if the nature of 
a problem is such that the problem would not significantly impact the 
safety or reliability of the airplane system involved. A relevant 
problem is a problem with an ETOPS group 1 significant system that has 
or could result in an IFSD or diversion. The applicant must include in 
this assessment relevant problems of supplier-provided ETOPS group 1 
significant systems and similar or identical equipment used on airplanes 
built by other manufacturers to the extent such information is 
reasonably available.
    (b) Propulsion system design. (1) The engine used in the applicant's 
airplane design must be approved as eligible for Early ETOPS in 
accordance with Sec. 33.201 of this chapter.
    (2) The applicant must design the propulsion system to preclude 
failures or malfunctions that could result in an IFSD. The applicant 
must show compliance with this requirement by analysis, test, in-service 
experience on other airplanes, or other means acceptable to the FAA. If 
analysis is used, the applicant must show that the propulsion system 
design will minimize failures and malfunctions with the objective of 
achieving the following IFSD rates:
    (i) An IFSD rate of 0.02 or less per 1,000 world-fleet engine-hours 
for type design approval up to and including 180 minutes.
    (ii) An IFSD rate of 0.01 or less per 1,000 world-fleet engine-hours 
for type design approval beyond 180 minutes.
    (c) Maintenance and operational procedures. The applicant must 
validate all maintenance and operational procedures for ETOPS 
significant systems. The applicant must identify, track, and resolve any 
problems found during the validation in accordance with the problem 
tracking and resolution system specified in section K25.2.2(h) of this 
appendix.
    (d) Propulsion system validation test. (1) The installed engine 
configuration for which approval is being sought must comply with Sec. 
33.201(c) of this chapter. The test engine must be configured with a 
complete airplane nacelle package, including engine-mounted equipment, 
except for any configuration differences necessary to accommodate test 
stand interfaces with the engine nacelle package. At the conclusion of 
the test, the propulsion system must be--
    (i) Visually inspected according to the applicant's on-wing 
inspection recommendations and limits; and
    (ii) Completely disassembled and the propulsion system hardware 
inspected to determine whether it meets the service limits specified in 
the Instructions for Continued Airworthiness submitted in compliance 
with Sec. 25.1529.
    (2) The applicant must identify, track, and resolve each cause or 
potential cause of IFSD, loss of thrust control, or other power loss 
encountered during this inspection in accordance with the problem 
tracking and resolution system specified in section K25.2.2 (h) of this 
appendix.
    (e) New technology testing. Technology new to the applicant, 
including substantially new manufacturing techniques, must be tested to 
substantiate its suitability for the airplane design.
    (f) APU validation test. If an APU is needed to comply with this 
appendix, one APU of the type to be certified with the airplane must be 
tested for 3,000 equivalent airplane operational cycles. Following 
completion of the test, the APU must be disassembled and inspected. The 
applicant must identify, track, and resolve each cause or potential 
cause of an inability to start or operate the APU in flight as intended 
in accordance with the problem tracking and resolution system specified 
in section K25.2.2(h) of this appendix.
    (g) Airplane demonstration. For each airplane-engine combination to 
be approved for ETOPS, the applicant must flight test at least one 
airplane to demonstrate that the airplane, and its components and 
equipment are capable of functioning properly during ETOPS flights and 
diversions of the longest duration for which the applicant seeks 
approval. This flight testing may be performed in conjunction with, but 
may not substitute for the flight testing required by Sec. 21.35(b)(2) 
of this chapter.
    (1) The airplane demonstration flight test program must include:
    (i) Flights simulating actual ETOPS, including flight at normal 
cruise altitude, step climbs, and, if applicable, APU operation.
    (ii) Maximum duration flights with maximum duration diversions.
    (iii) Maximum duration engine-inoperative diversions distributed 
among the engines installed on the airplanes used for the airplane 
demonstration flight test program. At least two one-engine-inoperative 
diversions must be conducted at maximum continuous thrust or power using 
the same engine.
    (iv) Flights under non-normal conditions to demonstrate the 
flightcrew's ability to safely conduct an ETOPS diversion with worst-
case ETOPS significant system failures or malfunctions that could occur 
in service.
    (v) Diversions to airports that represent airports of the types used 
for ETOPS diversions.

[[Page 624]]

    (vi) Repeated exposure to humid and inclement weather on the ground 
followed by a long-duration flight at normal cruise altitude.
    (2) The airplane demonstration flight test program must validate the 
adequacy of the airplane's flying qualities and performance, and the 
flightcrew's ability to safely conduct an ETOPS diversion under the 
conditions specified in section K25.2.2(g)(1) of this appendix.
    (3) During the airplane demonstration flight test program, each test 
airplane must be operated and maintained using the applicant's 
recommended operating and maintenance procedures.
    (4) At the completion of the airplane demonstration flight test 
program, each ETOPS significant system must undergo an on-wing 
inspection or test in accordance with the tasks defined in the proposed 
Instructions for Continued Airworthiness to establish its condition for 
continued safe operation. Each engine must also undergo a gas path 
inspection. These inspections must be conducted in a manner to identify 
abnormal conditions that could result in an IFSD or diversion. The 
applicant must identify, track and resolve any abnormal conditions in 
accordance with the problem tracking and resolution system specified in 
section K25.2.2(h) of this appendix.
    (h) Problem tracking and resolution system. (1) The applicant must 
establish and maintain a problem tracking and resolution system. The 
system must:
    (i) Contain a process for prompt reporting to the responsible FAA 
aircraft certification office of each occurrence reportable under Sec. 
21.4(a)(6) encountered during the phases of airplane and engine 
development used to assess Early ETOPS eligibility.
    (ii) Contain a process for notifying the responsible FAA aircraft 
certification office of each proposed corrective action that the 
applicant determines necessary for each problem identified from the 
occurrences reported under section K25.2.2. (h)(1)(i) of this appendix. 
The timing of the notification must permit appropriate FAA review before 
taking the proposed corrective action.
    (2) If the applicant is seeking ETOPS type design approval of a 
change to an airplane-engine combination previously approved for ETOPS, 
the problem tracking and resolution system need only address those 
problems specified in the following table, provided the applicant 
obtains prior authorization from the FAA:

------------------------------------------------------------------------
                                          Then the Problem Tracking and
  If the change does not require a new    Resolution System must address
  airplane type certificiate and . . .                . . .
------------------------------------------------------------------------
(i) Requires a new engine type           All problems applicable to the
 certificate.                             new engine installation, and
                                          for the remainder of the
                                          airplane, problems in changed
                                          systems only.
(ii) Does not require a new engine type  Problems in changed systems
 certificate.                             only.
------------------------------------------------------------------------

    (i) Acceptance criteria. The type and frequency of failures and 
malfunctions on ETOPS significant systems that occur during the airplane 
flight test program and the airplane demonstration flight test program 
specified in section K25.2.2(g) of this appendix must be consistent with 
the type and frequency of failures and malfunctions that would be 
expected to occur on currently certificated airplanes approved for 
ETOPS.
    K25.2.3. Combined service experience and Early ETOPS method.
    An applicant for ETOPS type design approval using the combined 
service experience and Early ETOPS method must comply with the following 
requirements.
    (a) A service experience requirement of not less than 15,000 engine-
hours for the world fleet of the candidate airplane-engine combination.
    (b) The Early ETOPS requirements of K25.2.2, except for the airplane 
demonstration specified in section K25.2.2(g) of this appendix; and
    (c) The flight test requirement of section K25.2.1(e) of this 
appendix.
    K25.3. Airplanes with more than two engines.
    An applicant for ETOPS type design approval of an airplane with more 
than two engines must use one of the methods described in section 
K25.3.1, K25.3.2, or K25.3.3 of this appendix.
    K25.3.1 Service experience method.
    An applicant for ETOPS type design approval using the service 
experience method must comply with section K25.3.1(a) of this appendix 
before conducting the airplane systems assessment specified in 
K25.3.1(b), and the flight test specified in section K25.3.1(c) of this 
appendix.
    (a) Service experience. The world fleet for the airplane-engine 
combination must accumulate a minimum of 250,000 engine-hours. The FAA 
may reduce this number of hours if the applicant identifies compensating 
factors that are acceptable to the FAA. The compensating factors may 
include experience on another airplane, but experience on the candidate 
airplane must make up a significant portion of the total required 
service experience.
    (b) Airplane systems assessment. The applicant must conduct an 
airplane systems assessment. The applicant must show that the

[[Page 625]]

airplane systems comply with the Sec. 25.1309(b) using available in-
service reliability data for ETOPS significant systems on the candidate 
airplane-engine combination. Each cause or potential cause of a relevant 
design, manufacturing, operational or maintenance problem occurring in 
service must have a corrective action or actions that are shown to be 
effective in preventing future occurrences. Each corrective action must 
be identified in the CMP document specified in section K25.1.6 of this 
appendix. A corrective action is not required if the problem would not 
significantly impact the safety or reliability of the airplane system 
involved. A relevant problem is a problem with an ETOPS group 1 
significant system that has or could result in an IFSD or diversion. The 
applicant must include in this assessment relevant problems with similar 
or identical equipment installed on other types of airplanes to the 
extent such information is reasonably available.
    (c) Airplane flight test. The applicant must conduct a flight test 
to validate the flightcrew's ability to safely conduct an ETOPS 
diversion with an inoperative engine and worst-case ETOPS significant 
system failures and malfunctions that could occur in service. The flight 
test must validate the airplane's flying qualities and performance with 
the demonstrated failures and malfunctions.
    K25.3.2 Early ETOPS method.
    An applicant for ETOPS type design approval using the Early ETOPS 
method must comply with the following requirements:
    (a) Maintenance and operational procedures. The applicant must 
validate all maintenance and operational procedures for ETOPS 
significant systems. The applicant must identify, track and resolve any 
problems found during the validation in accordance with the problem 
tracking and resolution system specified in section K25.3.2(e) of this 
appendix.
    (b) New technology testing. Technology new to the applicant, 
including substantially new manufacturing techniques, must be tested to 
substantiate its suitability for the airplane design.
    (c) APU validation test. If an APU is needed to comply with this 
appendix, one APU of the type to be certified with the airplane must be 
tested for 3,000 equivalent airplane operational cycles. Following 
completion of the test, the APU must be disassembled and inspected. The 
applicant must identify, track, and resolve each cause or potential 
cause of an inability to start or operate the APU in flight as intended 
in accordance with the problem tracking and resolution system specified 
in section K25.3.2(e) of this appendix.
    (d) Airplane demonstration. For each airplane-engine combination to 
be approved for ETOPS, the applicant must flight test at least one 
airplane to demonstrate that the airplane, and its components and 
equipment are capable of functioning properly during ETOPS flights and 
diversions of the longest duration for which the applicant seeks 
approval. This flight testing may be performed in conjunction with, but 
may not substitute for the flight testing required by Sec. 21.35(b)(2).
    (1) The airplane demonstration flight test program must include:
    (i) Flights simulating actual ETOPS including flight at normal 
cruise altitude, step climbs, and, if applicable, APU operation.
    (ii) Maximum duration flights with maximum duration diversions.
    (iii) Maximum duration engine-inoperative diversions distributed 
among the engines installed on the airplanes used for the airplane 
demonstration flight test program. At least two one engine-inoperative 
diversions must be conducted at maximum continuous thrust or power using 
the same engine.
    (iv) Flights under non-normal conditions to validate the 
flightcrew's ability to safely conduct an ETOPS diversion with worst-
case ETOPS significant system failures or malfunctions that could occur 
in service.
    (v) Diversions to airports that represent airports of the types used 
for ETOPS diversions.
    (vi) Repeated exposure to humid and inclement weather on the ground 
followed by a long duration flight at normal cruise altitude.
    (2) The airplane demonstration flight test program must validate the 
adequacy of the airplane's flying qualities and performance, and the 
flightcrew's ability to safely conduct an ETOPS diversion under the 
conditions specified in section K25.3.2(d)(1) of this appendix.
    (3) During the airplane demonstration flight test program, each test 
airplane must be operated and maintained using the applicant's 
recommended operating and maintenance procedures.
    (4) At the completion of the airplane demonstration, each ETOPS 
significant system must undergo an on-wing inspection or test in 
accordance with the tasks defined in the proposed Instructions for 
Continued Airworthiness to establish its condition for continued safe 
operation. Each engine must also undergo a gas path inspection. These 
inspections must be conducted in a manner to identify abnormal 
conditions that could result in an IFSD or diversion. The applicant must 
identify, track and resolve any abnormal conditions in accordance with 
the problem tracking and resolution system specified in section 
K25.3.2(e) of this appendix.
    (e) Problem tracking and resolution system. (1) The applicant must 
establish and maintain a problem tracking and resolution system. The 
system must:
    (i) Contain a process for prompt reporting to the responsible FAA 
aircraft certification office of each occurrence reportable under

[[Page 626]]

Sec. 21.4(a)(6) encountered during the phases of airplane and engine 
development used to assess Early ETOPS eligibility.
    (ii) Contain a process for notifying the responsible FAA aircraft 
certification office of each proposed corrective action that the 
applicant determines necessary for each problem identified from the 
occurrences reported under section K25.3.2(h)(1)(i) of this appendix. 
The timing of the notification must permit appropriate FAA review before 
taking the proposed corrective action.
    (2) If the applicant is seeking ETOPS type design approval of a 
change to an airplane-engine combination previously approved for ETOPS, 
the problem tracking and resolution system need only address those 
problems specified in the following table, provided the applicant 
obtains prior authorization from the FAA:

------------------------------------------------------------------------
                                          Then the Problem Tracking and
  If the change does not require a new    Resolution System must address
  airplane type certificate and . . .                 . . .
------------------------------------------------------------------------
(i) Requires a new engine type           All problems applicable to the
 certificate.                             new engine installation, and
                                          for the remainder of the
                                          airplane, problems in changed
                                          systems only.
(ii) Does not require a new engine type  Problems in changed systems
 certificate.                             only.
------------------------------------------------------------------------

    (f) Acceptance criteria. The type and frequency of failures and 
malfunctions on ETOPS significant systems that occur during the airplane 
flight test program and the airplane demonstration flight test program 
specified in section K25.3.2(d) of this appendix must be consistent with 
the type and frequency of failures and malfunctions that would be 
expected to occur on currently certificated airplanes approved for 
ETOPS.
    K25.3.3 Combined service experience and Early ETOPS method.
    An applicant for ETOPS type design approval using the Early ETOPS 
method must comply with the following requirements:
    (a) A service experience requirement of less than 15,000 engine-
hours for the world fleet of the candidate airplane-engine combination;
    (b) The Early ETOPS requirements of section K25.3.2 of this 
appendix, except for the airplane demonstration specified in section 
K25.3.2(d) of this appendix; and
    (c) The flight test requirement of section K25.3.1(c) of this 
appendix.

[Doc. No. FAA-2002-6717, 72 FR 1873, Jan. 16, 2007]



 Sec. Appendix L to Part 25--HIRF Environments and Equipment HIRF Test 
                                 Levels

    This appendix specifies the HIRF environments and equipment HIRF 
test levels for electrical and electronic systems under Sec. 25.1317. 
The field strength values for the HIRF environments and equipment HIRF 
test levels are expressed in root-mean-square units measured during the 
peak of the modulation cycle.
    (a) HIRF environment I is specified in the following table:

                      Table I.--HIRF Environment I
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-2 MHz......................................         50         50
2 MHz-30 MHz......................................        100        100
30 MHz-100 MHz....................................         50         50
100 MHz-400 MHz...................................        100        100
400 MHz-700 MHz...................................        700         50
700 MHz-1 GHz.....................................        700        100
1 GHz-2 GHz.......................................      2,000        200
2 GHz-6 GHz.......................................      3,000        200
6 GHz-8 GHz.......................................      1,000        200
8 GHz-12 GHz......................................      3,000        300
12 GHz-18 GHz.....................................      2,000        200
18 GHz-40 GHz.....................................        600        200
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (b) HIRF environment II is specified in the following table:

                      Table II.-HIRF Environment II
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-500 kHz....................................         20         20
500 kHz-2 MHz.....................................         30         30
2 MHz-30 MHz......................................        100        100
30 MHz-100 MHz....................................         10         10
100 MHz-200 MHz...................................         30         10
200 MHz-400 MHz...................................         10         10
400 MHz-1 GHz.....................................        700         40
1 GHz-2 GHz.......................................      1,300        160
2 GHz-4 GHz.......................................      3,000        120
4 GHz-6 GHz.......................................      3,000        160
6 GHz-8 GHz.......................................        400        170
8 GHz-12 GHz......................................      1,230        230
12 GHz-18 GHz.....................................        730        190
18 GHz-40 GHz.....................................        600        150
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (c) Equipment HIRF Test Level 1.
    (1) From 10 kilohertz (kHz) to 400 megahertz (MHz), use conducted 
susceptibility tests with continuous wave (CW) and 1 kHz

[[Page 627]]

square wave modulation with 90 percent depth or greater. The conducted 
susceptibility current must start at a minimum of 0.6 milliamperes (mA) 
at 10 kHz, increasing 20 decibels (dB) per frequency decade to a minimum 
of 30 mA at 500 kHz.
    (2) From 500 kHz to 40 MHz, the conducted susceptibility current 
must be at least 30 mA.
    (3) From 40 MHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 30 mA at 40 MHz, decreasing 20 dB per frequency 
decade to a minimum of 3 mA at 400 MHz.
    (4) From 100 MHz to 400 MHz, use radiated susceptibility tests at a 
minimum of 20 volts per meter (V/m) peak with CW and 1 kHz square wave 
modulation with 90 percent depth or greater.
    (5) From 400 MHz to 8 gigahertz (GHz), use radiated susceptibility 
tests at a minimum of 150 V/m peak with pulse modulation of 4 percent 
duty cycle with a 1 kHz pulse repetition frequency. This signal must be 
switched on and off at a rate of 1 Hz with a duty cycle of 50 percent.
    (d) Equipment HIRF Test Level 2. Equipment HIRF test level 2 is HIRF 
environment II in table II of this appendix reduced by acceptable 
aircraft transfer function and attenuation curves. Testing must cover 
the frequency band of 10 kHz to 8 GHz.
    (e) Equipment HIRF Test Level 3.
    (1) From 10 kHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 0.15 mA at 10 kHz, increasing 20 dB per 
frequency decade to a minimum of 7.5 mA at 500 kHz.
    (2) From 500 kHz to 40 MHz, use conducted susceptibility tests at a 
minimum of 7.5 mA.
    (3) From 40 MHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 7.5 mA at 40 MHz, decreasing 20 dB per 
frequency decade to a minimum of 0.75 mA at 400 MHz.
    (4) From 100 MHz to 8 GHz, use radiated susceptibility tests at a 
minimum of 5 V/m.

[Doc. No. FAA-2006-23657, 72 FR 44026, Aug. 6, 2007]



  Sec. Appendix M to Part 25--Fuel Tank System Flammability Reduction 
                                  Means

    M25.1 Fuel tank flammability exposure requirements.
    (a) The Fleet Average Flammability Exposure of each fuel tank, as 
determined in accordance with Appendix N of this part, may not exceed 3 
percent of the Flammability Exposure Evaluation Time (FEET), as defined 
in Appendix N of this part. As a portion of this 3 percent, if 
flammability reduction means (FRM) are used, each of the following time 
periods may not exceed 1.8 percent of the FEET:
    (1) When any FRM is operational but the fuel tank is not inert and 
the tank is flammable; and
    (2) When any FRM is inoperative and the tank is flammable.
    (b) The Fleet Average Flammability Exposure, as defined in Appendix 
N of this part, of each fuel tank may not exceed 3 percent of the 
portion of the FEET occurring during either ground or takeoff/climb 
phases of flight during warm days. The analysis must consider the 
following conditions.
    (1) The analysis must use the subset of those flights that begin 
with a sea level ground ambient temperature of 80 [deg]F (standard day 
plus 21 [deg]F atmosphere) or above, from the flammability exposure 
analysis done for overall performance.
    (2) For the ground and takeoff/climb phases of flight, the average 
flammability exposure must be calculated by dividing the time during the 
specific flight phase the fuel tank is flammable by the total time of 
the specific flight phase.
    (3) Compliance with this paragraph may be shown using only those 
flights for which the airplane is dispatched with the flammability 
reduction means operational.
    M25.2 Showing compliance.
    (a) The applicant must provide data from analysis, ground testing, 
and flight testing, or any combination of these, that:
    (1) Validate the parameters used in the analysis required by 
paragraph M25.1 of this appendix;
    (2) Substantiate that the FRM is effective at limiting flammability 
exposure in all compartments of each tank for which the FRM is used to 
show compliance with paragraph M25.1 of this appendix; and
    (3) Describe the circumstances under which the FRM would not be 
operated during each phase of flight.
    (b) The applicant must validate that the FRM meets the requirements 
of paragraph M25.1 of this appendix with any airplane or engine 
configuration affecting the performance of the FRM for which approval is 
sought.
    M25.3 Reliability indications and maintenance access.
    (a) Reliability indications must be provided to identify failures of 
the FRM that would otherwise be latent and whose identification is 
necessary to ensure the fuel tank with an FRM meets the fleet average 
flammability exposure requirements listed in paragraph M25.1 of this 
appendix, including when the FRM is inoperative.
    (b) Sufficient accessibility to FRM reliability indications must be 
provided for maintenance personnel or the flightcrew.
    (c) The access doors and panels to the fuel tanks with FRMs 
(including any tanks that

[[Page 628]]

communicate with a tank via a vent system), and to any other confined 
spaces or enclosed areas that could contain hazardous atmosphere under 
normal conditions or failure conditions, must be permanently stenciled, 
marked, or placarded to warn maintenance personnel of the possible 
presence of a potentially hazardous atmosphere.
    M25.4 Airworthiness limitations and procedures.
    (a) If FRM is used to comply with paragraph M25.1 of this appendix, 
Airworthiness Limitations must be identified for all maintenance or 
inspection tasks required to identify failures of components within the 
FRM that are needed to meet paragraph M25.1 of this appendix.
    (b) Maintenance procedures must be developed to identify any hazards 
to be considered during maintenance of the FRM. These procedures must be 
included in the instructions for continued airworthiness (ICA).
    M25.5 Reliability reporting.
    The effects of airplane component failures on FRM reliability must 
be assessed on an on-going basis. The applicant/holder must do the 
following:
    (a) Demonstrate effective means to ensure collection of FRM 
reliability data. The means must provide data affecting FRM reliability, 
such as component failures.
    (b) Unless alternative reporting procedures are approved by the FAA 
Oversight Office, as defined in part 26 of this subchapter, provide a 
report to the FAA every six months for the first five years after 
service introduction. After that period, continued reporting every six 
months may be replaced with other reliability tracking methods found 
acceptable to the FAA or eliminated if it is established that the 
reliability of the FRM meets, and will continue to meet, the exposure 
requirements of paragraph M25.1 of this appendix.
    (c) Develop service instructions or revise the applicable airplane 
manual, according to a schedule approved by the FAA Oversight Office, as 
defined in part 26 of this subchapter, to correct any failures of the 
FRM that occur in service that could increase any fuel tank's Fleet 
Average Flammability Exposure to more than that required by paragraph 
M25.1 of this appendix.

[Doc. No. FAA-2005-22997, 73 FR 42494, July 21, 2008]



    Sec. Appendix N TO Part 25--Fuel Tank Flammability Exposure and 
                          Reliability Analysis

    N25.1 General.
    (a) This appendix specifies the requirements for conducting fuel 
tank fleet average flammability exposure analyses required to meet Sec. 
25.981(b) and Appendix M of this part. For fuel tanks installed in 
aluminum wings, a qualitative assessment is sufficient if it 
substantiates that the tank is a conventional unheated wing tank.
    (b) This appendix defines parameters affecting fuel tank 
flammability that must be used in performing the analysis. These include 
parameters that affect all airplanes within the fleet, such as a 
statistical distribution of ambient temperature, fuel flash point, 
flight lengths, and airplane descent rate. Demonstration of compliance 
also requires application of factors specific to the airplane model 
being evaluated. Factors that need to be included are maximum range, 
cruise mach number, typical altitude where the airplane begins initial 
cruise phase of flight, fuel temperature during both ground and flight 
times, and the performance of a flammability reduction means (FRM) if 
installed.
    (c) The following definitions, input variables, and data tables must 
be used in the program to determine fleet average flammability exposure 
for a specific airplane model.
    N25.2 Definitions.
    (a) Bulk Average Fuel Temperature means the average fuel temperature 
within the fuel tank or different sections of the tank if the tank is 
subdivided by baffles or compartments.
    (b) Flammability Exposure Evaluation Time (FEET). The time from the 
start of preparing the airplane for flight, through the flight and 
landing, until all payload is unloaded, and all passengers and crew have 
disembarked. In the Monte Carlo program, the flight time is randomly 
selected from the Flight Length Distribution (Table 2), the pre-flight 
times are provided as a function of the flight time, and the post-flight 
time is a constant 30 minutes.
    (c) Flammable. With respect to a fluid or gas, flammable means 
susceptible to igniting readily or to exploding (14 CFR Part 1, 
Definitions). A non-flammable ullage is one where the fuel-air vapor is 
too lean or too rich to burn or is inert as defined below. For the 
purposes of this appendix, a fuel tank that is not inert is considered 
flammable when the bulk average fuel temperature within the tank is 
within the flammable range for the fuel type being used. For any fuel 
tank that is subdivided into sections by baffles or compartments, the 
tank is considered flammable when the bulk average fuel temperature 
within any section of the tank, that is not inert, is within the 
flammable range for the fuel type being used.
    (d) Flash Point. The flash point of a flammable fluid means the 
lowest temperature at which the application of a flame to a heated 
sample causes the vapor to ignite momentarily, or ``flash.'' Table 1 of 
this appendix provides the flash point for the standard fuel to be used 
in the analysis.
    (e) Fleet average flammability exposure is the percentage of the 
flammability exposure evaluation time (FEET) each fuel tank

[[Page 629]]

ullage is flammable for a fleet of an airplane type operating over the 
range of flight lengths in a world-wide range of environmental 
conditions and fuel properties as defined in this appendix.
    (f) Gaussian Distribution is another name for the normal 
distribution, a symmetrical frequency distribution having a precise 
mathematical formula relating the mean and standard deviation of the 
samples. Gaussian distributions yield bell-shaped frequency curves 
having a preponderance of values around the mean with progressively 
fewer observations as the curve extends outward.
    (g) Hazardous atmosphere. An atmosphere that may expose maintenance 
personnel, passengers or flight crew to the risk of death, 
incapacitation, impairment of ability to self-rescue (that is, escape 
unaided from a confined space), injury, or acute illness.
    (h) Inert. For the purpose of this appendix, the tank is considered 
inert when the bulk average oxygen concentration within each compartment 
of the tank is 12 percent or less from sea level up to 10,000 feet 
altitude, then linearly increasing from 12 percent at 10,000 feet to 
14.5 percent at 40,000 feet altitude, and extrapolated linearly above 
that altitude.
    (i) Inerting. A process where a noncombustible gas is introduced 
into the ullage of a fuel tank so that the ullage becomes non-flammable.
    (j) Monte Carlo Analysis. The analytical method that is specified in 
this appendix as the compliance means for assessing the fleet average 
flammability exposure time for a fuel tank.
    (k) Oxygen evolution occurs when oxygen dissolved in the fuel is 
released into the ullage as the pressure and temperature in the fuel 
tank are reduced.
    (l) Standard deviation is a statistical measure of the dispersion or 
variation in a distribution, equal to the square root of the arithmetic 
mean of the squares of the deviations from the arithmetic means.
    (m) Transport Effects. For purposes of this appendix, transport 
effects are the change in fuel vapor concentration in a fuel tank caused 
by low fuel conditions and fuel condensation and vaporization.
    (n) Ullage. The volume within the fuel tank not occupied by liquid 
fuel.
    N25.3 Fuel tank flammability exposure analysis.
    (a) A flammability exposure analysis must be conducted for the fuel 
tank under evaluation to determine fleet average flammability exposure 
for the airplane and fuel types under evaluation. For fuel tanks that 
are subdivided by baffles or compartments, an analysis must be performed 
either for each section of the tank, or for the section of the tank 
having the highest flammability exposure. Consideration of transport 
effects is not allowed in the analysis. The analysis must be done in 
accordance with the methods and procedures set forth in the Fuel Tank 
Flammability Assessment Method User's Manual, dated May 2008, document 
number DOT/FAA/AR-05/8 (incorporated by reference, see Sec. 25.5). The 
parameters specified in sections N25.3(b) and (c) of this appendix must 
be used in the fuel tank flammability exposure ``Monte Carlo'' analysis.
    (b) The following parameters are defined in the Monte Carlo analysis 
and provided in paragraph N25.4 of this appendix:
    (1) Cruise Ambient Temperature, as defined in this appendix.
    (2) Ground Ambient Temperature, as defined in this appendix.
    (3) Fuel Flash Point, as defined in this appendix.
    (4) Flight Length Distribution, as defined in Table 2 of this 
appendix.
    (5) Airplane Climb and Descent Profiles, as defined in the Fuel Tank 
Flammability Assessment Method User's Manual, dated May 2008, document 
number DOT/FAA/AR-05/8 (incorporated by reference in Sec. 25.5).
    (c) Parameters that are specific to the particular airplane model 
under evaluation that must be provided as inputs to the Monte Carlo 
analysis are:
    (1) Airplane cruise altitude.
    (2) Fuel tank quantities. If fuel quantity affects fuel tank 
flammability, inputs to the Monte Carlo analysis must be provided that 
represent the actual fuel quantity within the fuel tank or compartment 
of the fuel tank throughout each of the flights being evaluated. Input 
values for this data must be obtained from ground and flight test data 
or the approved FAA fuel management procedures.
    (3) Airplane cruise mach number.
    (4) Airplane maximum range.
    (5) Fuel tank thermal characteristics. If fuel temperature affects 
fuel tank flammability, inputs to the Monte Carlo analysis must be 
provided that represent the actual bulk average fuel temperature within 
the fuel tank at each point in time throughout each of the flights being 
evaluated. For fuel tanks that are subdivided by baffles or 
compartments, bulk average fuel temperature inputs must be provided for 
each section of the tank. Input values for these data must be obtained 
from ground and flight test data or a thermal model of the tank that has 
been validated by ground and flight test data.
    (6) Maximum airplane operating temperature limit, as defined by any 
limitations in the airplane flight manual.
    (7) Airplane Utilization. The applicant must provide data supporting 
the number of flights per day and the number of hours per flight for the 
specific airplane model under evaluation. If there is no existing 
airplane

[[Page 630]]

fleet data to support the airplane being evaluated, the applicant must 
provide substantiation that the number of flights per day and the number 
of hours per flight for that airplane model is consistent with the 
existing fleet data they propose to use.
    (d) Fuel Tank FRM Model. If FRM is used, an FAA approved Monte Carlo 
program must be used to show compliance with the flammability 
requirements of Sec. 25.981 and Appendix M of this part. The program 
must determine the time periods during each flight phase when the fuel 
tank or compartment with the FRM would be flammable. The following 
factors must be considered in establishing these time periods:
    (1) Any time periods throughout the flammability exposure evaluation 
time and under the full range of expected operating conditions, when the 
FRM is operating properly but fails to maintain a non-flammable fuel 
tank because of the effects of the fuel tank vent system or other 
causes,
    (2) If dispatch with the system inoperative under the Master Minimum 
Equipment List (MMEL) is requested, the time period assumed in the 
reliability analysis (60 flight hours must be used for a 10-day MMEL 
dispatch limit unless an alternative period has been approved by the 
Administrator),
    (3) Frequency and duration of time periods of FRM inoperability, 
substantiated by test or analysis acceptable to the FAA, caused by 
latent or known failures, including airplane system shut-downs and 
failures that could cause the FRM to shut down or become inoperative.
    (4) Effects of failures of the FRM that could increase the 
flammability exposure of the fuel tank.
    (5) If an FRM is used that is affected by oxygen concentrations in 
the fuel tank, the time periods when oxygen evolution from the fuel 
results in the fuel tank or compartment exceeding the inert level. The 
applicant must include any times when oxygen evolution from the fuel in 
the tank or compartment under evaluation would result in a flammable 
fuel tank. The oxygen evolution rate that must be used is defined in the 
Fuel Tank Flammability Assessment Method User's Manual, dated May 2008, 
document number DOT/FAA/AR-05/8 (incorporated by reference in Sec. 
25.5).
    (6) If an inerting system FRM is used, the effects of any air that 
may enter the fuel tank following the last flight of the day due to 
changes in ambient temperature, as defined in Table 4, during a 12-hour 
overnight period.
    (e) The applicant must submit to the FAA Oversight Office for 
approval the fuel tank flammability analysis, including the airplane-
specific parameters identified under paragraph N25.3(c) of this appendix 
and any deviations from the parameters identified in paragraph N25.3(b) 
of this appendix that affect flammability exposure, substantiating data, 
and any airworthiness limitations and other conditions assumed in the 
analysis.
    N25.4 Variables and data tables.
    The following data must be used when conducting a flammability 
exposure analysis to determine the fleet average flammability exposure. 
Variables used to calculate fleet flammability exposure must include 
atmospheric ambient temperatures, flight length, flammability exposure 
evaluation time, fuel flash point, thermal characteristics of the fuel 
tank, overnight temperature drop, and oxygen evolution from the fuel 
into the ullage.
    (a) Atmospheric Ambient Temperatures and Fuel Properties.
    (1) In order to predict flammability exposure during a given flight, 
the variation of ground ambient temperatures, cruise ambient 
temperatures, and a method to compute the transition from ground to 
cruise and back again must be used. The variation of the ground and 
cruise ambient temperatures and the flash point of the fuel is defined 
by a Gaussian curve, given by the 50 percent value and a 1-standard deviation value.
    (2) Ambient Temperature: Under the program, the ground and cruise 
ambient temperatures are linked by a set of assumptions on the 
atmosphere. The temperature varies with altitude following the 
International Standard Atmosphere (ISA) rate of change from the ground 
ambient temperature until the cruise temperature for the flight is 
reached. Above this altitude, the ambient temperature is fixed at the 
cruise ambient temperature. This results in a variation in the upper 
atmospheric temperature. For cold days, an inversion is applied up to 
10,000 feet, and then the ISA rate of change is used.
    (3) Fuel properties:
    (i) For Jet A fuel, the variation of flash point of the fuel is 
defined by a Gaussian curve, given by the 50 percent value and a 1-standard deviation, as shown in Table 1 of this 
appendix.
    (ii) The flammability envelope of the fuel that must be used for the 
flammability exposure analysis is a function of the flash point of the 
fuel selected by the Monte Carlo for a given flight. The flammability 
envelope for the fuel is defined by the upper flammability limit (UFL) 
and lower flammability limit (LFL) as follows:
    (A) LFL at sea level = flash point temperature of the fuel at sea 
level minus 10 [deg]F. LFL decreases from sea level value with 
increasing altitude at a rate of 1 [deg]F per 808 feet.
    (B) UFL at sea level = flash point temperature of the fuel at sea 
level plus 63.5 [deg]F. UFL decreases from the sea level value with 
increasing altitude at a rate of 1 [deg]F per 512 feet.
    (4) For each flight analyzed, a separate random number must be 
generated for each of the three parameters (ground ambient

[[Page 631]]

temperature, cruise ambient temperature, and fuel flash point) using the 
Gaussian distribution defined in Table 1 of this appendix.

Table 1.--Gaussian Distribution for Ground Ambient Temperature, Cruise Ambient Temperature, and Fuel Flash Point
----------------------------------------------------------------------------------------------------------------
                                                                            Temperature in deg F
                                                           -----------------------------------------------------
                         Parameter                           Ground ambient    Cruise ambient   Fuel flash point
                                                               temperature       temperature          (FP)
----------------------------------------------------------------------------------------------------------------
Mean Temp.................................................             59.95               -70               120
Neg 1 std dev.............................................             20.14                 8                 8
Pos 1 std dev.............................................             17.28                 8                 8
----------------------------------------------------------------------------------------------------------------

    (b) The Flight Length Distribution defined in Table 2 must be used 
in the Monte Carlo analysis.

                                      Table 2.--Flight Length Distribution
----------------------------------------------------------------------------------------------------------------
Flight length (NM)                           Airplane maximum range--nautical miles (NM)
----------------------------------------------------------------------------------------------------------------
  From       To       1000      2000      3000      4000      5000     6000     7000     8000     9000    10000
----------------------------------------------------------------------------------------------------------------
          ........                      Distribution of flight lengths (percentage of total)
----------------------------------------------------------------------------------------------------------------
      0       200      11.7       7.5       6.2       5.5       4.7      4.0      3.4      3.0      2.6      2.3
    200       400      27.3      19.9      17.0      15.2      13.2     11.4      9.7      8.5      7.5      6.7
    400       600      46.3      40.0      35.7      32.6      28.5     24.9     21.2     18.7     16.4     14.8
    600       800      10.3      11.6      11.0      10.2       9.1      8.0      6.9      6.1      5.4      4.8
    800      1000       4.4       8.5       8.6       8.2       7.4      6.6      5.7      5.0      4.5      4.0
   1000      1200       0.0       4.8       5.3       5.3       4.8      4.3      3.8      3.3      3.0      2.7
   1200      1400       0.0       3.6       4.4       4.5       4.2      3.8      3.3      3.0      2.7      2.4
   1400      1600       0.0       2.2       3.3       3.5       3.3      3.1      2.7      2.4      2.2      2.0
   1600      1800       0.0       1.2       2.3       2.6       2.5      2.4      2.1      1.9      1.7      1.6
   1800      2000       0.0       0.7       2.2       2.6       2.6      2.5      2.2      2.0      1.8      1.7
   2000      2200       0.0       0.0       1.6       2.1       2.2      2.1      1.9      1.7      1.6      1.4
   2200      2400       0.0       0.0       1.1       1.6       1.7      1.7      1.6      1.4      1.3      1.2
   2400      2600       0.0       0.0       0.7       1.2       1.4      1.4      1.3      1.2      1.1      1.0
   2600      2800       0.0       0.0       0.4       0.9       1.0      1.1      1.0      0.9      0.9      0.8
   2800      3000       0.0       0.0       0.2       0.6       0.7      0.8      0.7      0.7      0.6      0.6
   3000      3200       0.0       0.0       0.0       0.6       0.8      0.8      0.8      0.8      0.7      0.7
   3200      3400       0.0       0.0       0.0       0.7       1.1      1.2      1.2      1.1      1.1      1.0
   3400      3600       0.0       0.0       0.0       0.7       1.3      1.6      1.6      1.5      1.5      1.4
   3600      3800       0.0       0.0       0.0       0.9       2.2      2.7      2.8      2.7      2.6      2.5
   3800      4000       0.0       0.0       0.0       0.5       2.0      2.6      2.8      2.8      2.7      2.6
   4000      4200       0.0       0.0       0.0       0.0       2.1      3.0      3.2      3.3      3.2      3.1
   4200      4400       0.0       0.0       0.0       0.0       1.4      2.2      2.5      2.6      2.6      2.5
   4400      4600       0.0       0.0       0.0       0.0       1.0      2.0      2.3      2.5      2.5      2.4
   4600      4800       0.0       0.0       0.0       0.0       0.6      1.5      1.8      2.0      2.0      2.0
   4800      5000       0.0       0.0       0.0       0.0       0.2      1.0      1.4      1.5      1.6      1.5
   5000      5200       0.0       0.0       0.0       0.0       0.0      0.8      1.1      1.3      1.3      1.3
   5200      5400       0.0       0.0       0.0       0.0       0.0      0.8      1.2      1.5      1.6      1.6
   5400      5600       0.0       0.0       0.0       0.0       0.0      0.9      1.7      2.1      2.2      2.3
   5600      5800       0.0       0.0       0.0       0.0       0.0      0.6      1.6      2.2      2.4      2.5
   5800      6000       0.0       0.0       0.0       0.0       0.0      0.2      1.8      2.4      2.8      2.9
   6000      6200       0.0       0.0       0.0       0.0       0.0      0.0      1.7      2.6      3.1      3.3
   6200      6400       0.0       0.0       0.0       0.0       0.0      0.0      1.4      2.4      2.9      3.1
   6400      6600       0.0       0.0       0.0       0.0       0.0      0.0      0.9      1.8      2.2      2.5
   6600      6800       0.0       0.0       0.0       0.0       0.0      0.0      0.5      1.2      1.6      1.9
   6800      7000       0.0       0.0       0.0       0.0       0.0      0.0      0.2      0.8      1.1      1.3
   7000      7200       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.4      0.7      0.8
   7200      7400       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.3      0.5      0.7
   7400      7600       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.2      0.5      0.6
   7600      7800       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.1      0.5      0.7
   7800      8000       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.1      0.6      0.8
   8000      8200       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.5      0.8
   8200      8400       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.5      1.0
   8400      8600       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.6      1.3
   8600      8800       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.4      1.1
   8800      9000       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.2      0.8
   9000      9200       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.0      0.5

[[Page 632]]

 
   9200      9400       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.0      0.2
   9400      9600       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.0      0.1
   9600      9800       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.0      0.1
   9800     10000       0.0       0.0       0.0       0.0       0.0      0.0      0.0      0.0      0.0      0.1
----------------------------------------------------------------------------------------------------------------

    (c) Overnight Temperature Drop. For airplanes on which FRM is 
installed, the overnight temperature drop for this appendix is defined 
using:
    (1) A temperature at the beginning of the overnight period that 
equals the landing temperature of the previous flight that is a random 
value based on a Gaussian distribution; and
    (2) An overnight temperature drop that is a random value based on a 
Gaussian distribution.
    (3) For any flight that will end with an overnight ground period 
(one flight per day out of an average number of flights per day, 
depending on utilization of the particular airplane model being 
evaluated), the landing outside air temperature (OAT) is to be chosen as 
a random value from the following Gaussian curve:

                Table 3.--Landing Outside Air Temperature
------------------------------------------------------------------------
                                                        Landing outside
                      Parameter                         air temperature
                                                             [deg]F
------------------------------------------------------------------------
Mean Temperature.....................................              58.68
negative 1 std dev...................................              20.55
positive 1 std dev...................................              13.21
------------------------------------------------------------------------

    (4) The outside ambient air temperature (OAT) overnight temperature 
drop is to be chosen as a random value from the following Gaussian 
curve:

              Table 4.--Outside Air Temperature (OAT) Drop
------------------------------------------------------------------------
                                                             OAT drop
                        Parameter                           temperature
                                                              [deg]F
------------------------------------------------------------------------
Mean Temp...............................................            12.0
1 std dev...............................................             6.0
------------------------------------------------------------------------

    (d) Number of Simulated Flights Required in Analysis. In order for 
the Monte Carlo analysis to be valid for showing compliance with the 
fleet average and warm day flammability exposure requirements, the 
applicant must run the analysis for a minimum number of flights to 
ensure that the fleet average and warm day flammability exposure for the 
fuel tank under evaluation meets the applicable flammability limits 
defined in Table 5 of this appendix.

                  Table 5.--Flammability Exposure Limit
------------------------------------------------------------------------
                                         Maximum            Maximum
                                     acceptable Monte   acceptable Monte
                                      Carlo average      Carlo average
                                        fuel tank          fuel tank
Minimum number of flights in Monte     flammability       flammability
          Carlo analysis                 exposure           exposure
                                    (percent) to meet  (percent) to meet
                                        3 percent      7 percent part 26
                                       requirements       requirements
------------------------------------------------------------------------
10,000............................               2.91               6.79
100,000...........................               2.98               6.96
1,000,000.........................               3.00               7.00
------------------------------------------------------------------------


[Doc. No. FAA-2005-22997, 73 FR 42495, July 21, 2008]



   Sec. Appendix O TO Part 25--Supercooled Large Drop Icing Conditions

    This Appendix consists of two parts. Part I defines this Appendix as 
a description of supercooled large drop icing conditions in which the 
drop median volume diameter (MVD) is less than or greater than 40 
[micro]m, the maximum mean effective drop diameter (MED) of Appendix C 
of this part continuous maximum (stratiform clouds) icing conditions. 
For this Appendix, supercooled large drop icing conditions consist of 
freezing drizzle and freezing rain occurring in and/or below stratiform 
clouds. Part II defines ice accretions used to show compliance with the 
airplane performance and handling qualities requirements of subpart B of 
this part.

                           PART I--METEOROLOGY

    In this Appendix icing conditions are defined by the parameters of 
altitude, vertical and horizontal extent, temperature, liquid water 
content, and water mass distribution as a function of drop diameter 
distribution.
    (a) Freezing Drizzle (Conditions with spectra maximum drop diameters 
from 100[micro]m to 500 [micro]m):
    (1) Pressure altitude range: 0 to 22,000 feet MSL.
    (2) Maximum vertical extent: 12,000 feet.
    (3) Horizontal extent: Standard distance of 17.4 nautical miles.
    (4) Total liquid water content.

    Note: Liquid water content (LWC) in grams per cubic meter (g/m\3\) 
based on horizontal extent standard distance of 17.4 nautical miles.
    (5) Drop diameter distribution: Figure 2.

[[Page 633]]

    (6) Altitude and temperature envelope: Figure 3.
    (b) Freezing Rain (Conditions with spectra maximum drop diameters 
greater than 500 [micro]m):
    (1) Pressure altitude range: 0 to 12,000 ft MSL.
    (2) Maximum vertical extent: 7,000 ft.
    (3) Horizontal extent: Standard distance of 17.4 nautical miles.
    (4) Total liquid water content.

    Note: LWC in grams per cubic meter (g/m\3\) based on horizontal 
extent standard distance of 17.4 nautical miles.
    (5) Drop Diameter Distribution: Figure 5.
    (6) Altitude and temperature envelope: Figure 6.
    (c) Horizontal extent.
    The liquid water content for freezing drizzle and freezing rain 
conditions for horizontal extents other than the standard 17.4 nautical 
miles can be determined by the value of the liquid water content 
determined from Figure 1 or Figure 4, multiplied by the factor provided 
in Figure 7, which is defined by the following equation:

S = 1.266 - 0.213 log10(H)
Where:
S = Liquid Water Content Scale Factor (dimensionless) and
H = horizontal extent in nautical miles
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[[Page 635]]


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[[Page 636]]


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[[Page 637]]


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[[Page 638]]


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[[Page 639]]


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 PART II--AIRFRAME ICE ACCRETIONS FOR SHOWING COMPLIANCE WITH SUBPART B 
                              OF THIS PART

    (a) General.
    The most critical ice accretion in terms of airplane performance and 
handling qualities for each flight phase must be used to show compliance 
with the applicable airplane performance and handling qualities 
requirements for icing conditions contained in subpart B of this part. 
Applicants must demonstrate that the full range of atmospheric icing 
conditions specified in part I of this Appendix have been considered, 
including drop diameter distributions, liquid water content, and 
temperature appropriate to the flight conditions (for example, 
configuration, speed, angle of attack, and altitude).
    (1) For an airplane certified in accordance with Sec. 
25.1420(a)(1), the ice accretions for each flight phase are defined in 
part II, paragraph (b) of this Appendix.
    (2) For an airplane certified in accordance with Sec. 
25.1420(a)(2), the most critical ice accretion for each flight phase 
defined in part II, paragraphs (b) and (c) of this Appendix, must be 
used. For the ice accretions defined in part II, paragraph (c) of this 
Appendix, only the portion of part I of this Appendix in which the 
airplane is capable of operating safely must be considered.
    (3) For an airplane certified in accordance with Sec. 
25.1420(a)(3), the ice accretions for each flight phase are defined in 
part II, paragraph (c) of this Appendix.
    (b) Ice accretions for airplanes certified in accordance with Sec. 
25.1420(a)(1) or (2).
    (1) En route ice is the en route ice as defined by part II, 
paragraph (c)(3), of this Appendix, for an airplane certified in 
accordance with Sec. 25.1420(a)(2), or defined by part II, paragraph 
(a)(3), of Appendix C of this part, for an airplane certified in 
accordance with Sec. 25.1420(a)(1), plus:
    (i) Pre-detection ice as defined by part II, paragraph (b)(5), of 
this Appendix; and
    (ii) The ice accumulated during the transit of one cloud with a 
horizontal extent of 17.4 nautical miles in the most critical of the 
icing conditions defined in part I of this Appendix and one cloud with a 
horizontal extent of 17.4 nautical miles in the continuous

[[Page 640]]

maximum icing conditions defined in Appendix C of this part.
    (2) Holding ice is the holding ice defined by part II, paragraph 
(c)(4), of this Appendix, for an airplane certified in accordance with 
Sec. 25.1420(a)(2), or defined by part II, paragraph (a)(4), of 
Appendix C of this part, for an airplane certified in accordance with 
Sec. 25.1420(a)(1), plus:
    (i) Pre-detection ice as defined by part II, paragraph (b)(5), of 
this Appendix; and
    (ii) The ice accumulated during the transit of one cloud with a 17.4 
nautical miles horizontal extent in the most critical of the icing 
conditions defined in part I of this Appendix and one cloud with a 
horizontal extent of 17.4 nautical miles in the continuous maximum icing 
conditions defined in Appendix C of this part.
    (iii) Except the total exposure to holding ice conditions does not 
need to exceed 45 minutes.
    (3) Approach ice is the more critical of the holding ice defined by 
part II, paragraph (b)(2), of this Appendix, or the ice calculated in 
the applicable paragraphs (b)(3)(i) or (ii) of part II, of this 
Appendix:
    (i) For an airplane certified in accordance with Sec. 
25.1420(a)(2), the ice accumulated during descent from the maximum 
vertical extent of the icing conditions defined in part I of this 
Appendix to 2,000 feet above the landing surface in the cruise 
configuration, plus transition to the approach configuration, plus:
    (A) Pre-detection ice, as defined by part II, paragraph (b)(5), of 
this Appendix; and
    (B) The ice accumulated during the transit at 2,000 feet above the 
landing surface of one cloud with a horizontal extent of 17.4 nautical 
miles in the most critical of the icing conditions defined in part I of 
this Appendix and one cloud with a horizontal extent of 17.4 nautical 
miles in the continuous maximum icing conditions defined in Appendix C 
of this part.
    (ii) For an airplane certified in accordance with Sec. 
25.1420(a)(1), the ice accumulated during descent from the maximum 
vertical extent of the maximum continuous icing conditions defined in 
part I of Appendix C to 2,000 feet above the landing surface in the 
cruise configuration, plus transition to the approach configuration, 
plus:
    (A) Pre-detection ice, as defined by part II, paragraph (b)(5), of 
this Appendix; and
    (B) The ice accumulated during the transit at 2,000 feet above the 
landing surface of one cloud with a horizontal extent of 17.4 nautical 
miles in the most critical of the icing conditions defined in part I of 
this Appendix and one cloud with a horizontal extent of 17.4 nautical 
miles in the continuous maximum icing conditions defined in Appendix C 
of this part.
    (4) Landing ice is the more critical of the holding ice as defined 
by part II, paragraph (b)(2), of this Appendix, or the ice calculated in 
the applicable paragraphs (b)(4)(i) or (ii) of part II of this Appendix:
    (i) For an airplane certified in accordance with Sec. 
25.1420(a)(2), the ice accretion defined by part II, paragraph 
(c)(5)(i), of this Appendix, plus a descent from 2,000 feet above the 
landing surface to a height of 200 feet above the landing surface with a 
transition to the landing configuration in the icing conditions defined 
in part I of this Appendix, plus:
    (A) Pre-detection ice, as defined in part II, paragraph (b)(5), of 
this Appendix; and
    (B) The ice accumulated during an exit maneuver, beginning with the 
minimum climb gradient required by Sec. 25.119, from a height of 200 
feet above the landing surface through one cloud with a horizontal 
extent of 17.4 nautical miles in the most critical of the icing 
conditions defined in part I of this Appendix and one cloud with a 
horizontal extent of 17.4 nautical miles in the continuous maximum icing 
conditions defined in Appendix C of this part.
    (ii) For an airplane certified in accordance with Sec. 
25.1420(a)(1), the ice accumulated in the maximum continuous icing 
conditions defined in Appendix C of this part, during a descent from the 
maximum vertical extent of the icing conditions defined in Appendix C of 
this part, to 2,000 feet above the landing surface in the cruise 
configuration, plus transition to the approach configuration and flying 
for 15 minutes at 2,000 feet above the landing surface, plus a descent 
from 2,000 feet above the landing surface to a height of 200 feet above 
the landing surface with a transition to the landing configuration, 
plus:
    (A) Pre-detection ice, as described by part II, paragraph (b)(5), of 
this Appendix; and
    (B) The ice accumulated during an exit maneuver, beginning with the 
minimum climb gradient required by Sec. 25.119, from a height of 200 
feet above the landing surface through one cloud with a horizontal 
extent of 17.4 nautical miles in the most critical of the icing 
conditions defined in part I of this Appendix and one cloud with a 
horizontal extent of 17.4 nautical miles in the continuous maximum icing 
conditions defined in Appendix C of this part.
    (5) Pre-detection ice is the ice accretion before detection of 
flight conditions in this Appendix that require exiting per Sec. 
25.1420(a)(1) and (2). It is the pre-existing ice accretion that may 
exist from operating in icing conditions in which the airplane is 
approved to operate prior to encountering the icing conditions requiring 
an exit, plus the ice accumulated during the time needed to detect the 
icing conditions, followed by two minutes of further ice accumulation to 
take into account the time for the flightcrew to take action to exit the 
icing conditions, including coordination with air traffic control.

[[Page 641]]

    (i) For an airplane certified in accordance with Sec. 
25.1420(a)(1), the pre-existing ice accretion must be based on the icing 
conditions defined in Appendix C of this part.
    (ii) For an airplane certified in accordance with Sec. 
25.1420(a)(2), the pre-existing ice accretion must be based on the more 
critical of the icing conditions defined in Appendix C of this part, or 
the icing conditions defined in part I of this Appendix in which the 
airplane is capable of safely operating.
    (c) Ice accretions for airplanes certified in accordance with 
Sec. Sec. 25.1420(a)(2) or (3). For an airplane certified in accordance 
with Sec. 25.1420(a)(2), only the portion of the icing conditions of 
part I of this Appendix in which the airplane is capable of operating 
safely must be considered.
    (1) Takeoff ice is the most critical ice accretion on unprotected 
surfaces, and any ice accretion on the protected surfaces, occurring 
between the end of the takeoff distance and 400 feet above the takeoff 
surface, assuming accretion starts at the end of the takeoff distance in 
the icing conditions defined in part I of this Appendix.
    (2) Final takeoff ice is the most critical ice accretion on 
unprotected surfaces, and any ice accretion on the protected surfaces 
appropriate to normal ice protection system operation, between 400 feet 
and either 1,500 feet above the takeoff surface, or the height at which 
the transition from the takeoff to the en route configuration is 
completed and VFTO is reached, whichever is higher. Ice 
accretion is assumed to start at the end of the takeoff distance in the 
icing conditions defined in part I of this Appendix.
    (3) En route ice is the most critical ice accretion on the 
unprotected surfaces, and any ice accretion on the protected surfaces 
appropriate to normal ice protection system operation, during the en 
route flight phase in the icing conditions defined in part I of this 
Appendix.
    (4) Holding ice is the most critical ice accretion on the 
unprotected surfaces, and any ice accretion on the protected surfaces 
appropriate to normal ice protection system operation, resulting from 45 
minutes of flight within a cloud with a 17.4 nautical miles horizontal 
extent in the icing conditions defined in part I of this Appendix, 
during the holding phase of flight.
    (5) Approach ice is the ice accretion on the unprotected surfaces, 
and any ice accretion on the protected surfaces appropriate to normal 
ice protection system operation, resulting from the more critical of 
the:
    (i) Ice accumulated in the icing conditions defined in part I of 
this Appendix during a descent from the maximum vertical extent of the 
icing conditions defined in part I of this Appendix, to 2,000 feet above 
the landing surface in the cruise configuration, plus transition to the 
approach configuration and flying for 15 minutes at 2,000 feet above the 
landing surface; or
    (ii) Holding ice as defined by part II, paragraph (c)(4), of this 
Appendix.
    (6) Landing ice is the ice accretion on the unprotected surfaces, 
and any ice accretion on the protected surfaces appropriate to normal 
ice protection system operation, resulting from the more critical of 
the:
    (i) Ice accretion defined by part II, paragraph (c)(5)(i), of this 
Appendix, plus ice accumulated in the icing conditions defined in part I 
of this Appendix during a descent from 2,000 feet above the landing 
surface to a height of 200 feet above the landing surface with a 
transition to the landing configuration, followed by a go-around at the 
minimum climb gradient required by Sec. 25.119, from a height of 200 
feet above the landing surface to 2,000 feet above the landing surface, 
flying for 15 minutes at 2,000 feet above the landing surface in the 
approach configuration, and a descent to the landing surface (touchdown) 
in the landing configuration; or
    (ii) Holding ice as defined by part II, paragraph (c)(4), of this 
Appendix.
    (7) For both unprotected and protected parts, the ice accretion for 
the takeoff phase must be determined for the icing conditions defined in 
part I of this Appendix, using the following assumptions:
    (i) The airfoils, control surfaces, and, if applicable, propellers 
are free from frost, snow, or ice at the start of takeoff;
    (ii) The ice accretion starts at the end of the takeoff distance;
    (iii) The critical ratio of thrust/power-to-weight;
    (iv) Failure of the critical engine occurs at VEF; and
    (v) Crew activation of the ice protection system is in accordance 
with a normal operating procedure provided in the airplane flight 
manual, except that after beginning the takeoff roll, it must be assumed 
that the crew takes no action to activate the ice protection system 
until the airplane is at least 400 feet above the takeoff surface.
    (d) The ice accretion before the ice protection system has been 
activated and is performing its intended function is the critical ice 
accretion formed on the unprotected and normally protected surfaces 
before activation and effective operation of the ice protection system 
in the icing conditions defined in part I of this Appendix. This ice 
accretion only applies in showing compliance to Sec. Sec. 25.143(j) and 
25.207(h).
    (e) In order to reduce the number of ice accretions to be considered 
when demonstrating compliance with the requirements of Sec. 25.21(g), 
any of the ice accretions defined in this Appendix may be used for any 
other flight phase if it is shown to be at least

[[Page 642]]

as critical as the specific ice accretion defined for that flight phase. 
Configuration differences and their effects on ice accretions must be 
taken into account.
    (f) The ice accretion that has the most adverse effect on handling 
qualities may be used for airplane performance tests provided any 
difference in performance is conservatively taken into account.

[79 FR 65528, Nov. 4, 2014]

    Effective Date Note: At 79 FR 65528, Nov. 4, 2014, Part 25, Appendix 
O was added, effective Jan. 5, 2015.



PART 26_CONTINUED AIRWORTHINESS AND SAFETY IMPROVEMENTS FOR TRANSPORT
CATEGORY AIRPLANES--Table of Contents



                            Subpart A_General

Sec.
26.1 Purpose and scope.
26.3 Definitions.
26.5 Applicability table.

      Subpart B_Enhanced Airworthiness Program for Airplane Systems

26.11 Electrical wiring interconnection systems (EWIS) maintenance 
          program.

        Subpart C_Aging Airplane Safety_Widespread Fatigue Damage

26.21 Limit of validity.
26.23 Extended limit of validity.

                    Subpart D_Fuel Tank Flammability

26.31 Definitions.
26.33 Holders of type certificates: Fuel tank flammability.
26.35 Changes to type certificates affecting fuel tank flammability.
26.37 Pending type certification projects: Fuel tank flammability.
26.39 Newly produced airplanes: Fuel tank flammability.

 Subpart E_Aging Airplane Safety_Damage Tolerance Data for Repairs and 
                               Alterations

26.41 Definitions.
26.43 Holders of and applicants for type certificates--Repairs.
26.45 Holders of type certificates--Alterations and repairs to 
          alterations.
26.47 Holders of and applicants for a supplemental type certificate--
          Alterations and repairs to alterations.
26.49 Compliance plan.

    Authority: 49 U.S.C. 106(g), 40113, 44701, 44702 and 44704.

    Source: Docket No. FAA-2004-18379, 72 FR 63409, Nov. 8, 2007, unless 
otherwise noted.



                            Subpart A_General



Sec. 26.1  Purpose and scope.

    (a) This part establishes requirements for support of the continued 
airworthiness of and safety improvements for transport category 
airplanes. These requirements may include performing assessments, 
developing design changes, developing revisions to Instructions for 
Continued Airworthiness (ICA), and making necessary documentation 
available to affected persons. Requirements of this part that establish 
standards for design changes and revisions to the ICA are considered 
airworthiness requirements.
    (b) Except as provided in paragraph (c) of this section, this part 
applies to the following persons, as specified in each subpart of this 
part:
    (1) Holders of type certificates and supplemental type certificates.
    (2) Applicants for type certificates and supplemental type 
certificates and changes to those certificates (including service 
bulletins describing design changes).
    (3) Persons seeking design approval for airplane repairs, 
alterations, or modifications that may affect airworthiness.
    (4) Holders of type certificates and their licensees producing new 
airplanes.
    (c) An applicant for approval of a design change is not required to 
comply with any applicable airworthiness requirement of this part if the 
applicant elects or is required to comply with a corresponding amendment 
to part 25 of this chapter that is adopted concurrently or after that 
airworthiness requirement.
    (d) For the purposes of this part, the word ``type certificate'' 
does not include supplemental type certificates.



Sec. 26.3  Definitions.

    For the purposes of this part:
    FAA Oversight Office is the aircraft certification office or office 
of the Transport Airplane Directorate with oversight responsibility for 
the relevant type certificate, supplemental

[[Page 643]]

type certificate, or manufacturer, as determined by the Administrator.



Sec. 26.5  Applicability table.

    Table 1 of this section provides an overview of the applicability of 
this part. It provides guidance in identifying what sections apply to 
various types of entities. The specific applicability of each subpart 
and section is specified in the regulatory text.

                                     Table 1--Applicability of Part 26 Rules
----------------------------------------------------------------------------------------------------------------
                                                                Applicable sections
                                 -------------------------------------------------------------------------------
                                                           Subpart C
                                   Subpart B EAPAS/   widespread fatigue    Subpart D fuel     Subpart E damage
                                          FTS               damage         tank flammability    tolerance data
----------------------------------------------------------------------------------------------------------------
Effective date of rule..........  December 10, 2007.  January 14, 2011..  December 26, 2008.  January 11, 2008
Existing \1\ TC Holders.........  26.11.............  26.21.............  26.33.............  26.43, 26.45,
                                                                                               26.49
Pending \1\ TC Applicants.......  26.11.............  26.21.............  26.37.............  26.43, 26.45
Future \2\ TC applicants........  N/A...............  N/A...............  N/A...............  26.43
Existing \1\ STC Holders........  N/A...............  26.21.............  26.35.............  26.47, 26.49
Pending \1\ STC/ATC applicants..  26.11.............  26.21.............  26.35.............  26.45, 26.47,
                                                                                               26.49
Future \2\ STC/ATC applicants...  26.11.............  26.21.............  26.35.............  26.45, 26.47,
                                                                                               26.49
Manufacturers...................  N/A...............  N/A...............  26.39.............  N/A
----------------------------------------------------------------------------------------------------------------
\1\ As of the effective date of the identified rule.
\2\ Application made after the effective date of the identified rule.


[Doc. No. FAA-2006-24281, 75 FR 69782, Nov. 15, 2010]



      Subpart B_Enhanced Airworthiness Program for Airplane Systems



Sec. 26.11  Electrical wiring interconnection systems (EWIS)
maintenance program.

    (a) Except as provided in paragraph (g) of this section, this 
section applies to transport category, turbine-powered airplanes with a 
type certificate issued after January 1, 1958, that, as a result of the 
original certification, or later increase in capacity, have--
    (1) A maximum type-certificated passenger capacity of 30 or more or
    (2) A maximum payload capacity of 7,500 pounds or more.
    (b) Holders of, and applicants for, type certificates, as identified 
in paragraph (d) of this section must develop Instructions for Continued 
Airworthiness (ICA) for the representative airplane's EWIS in accordance 
with part 25, Appendix H paragraphs H25.5(a)(1) and (b) of this 
subchapter in effect on December 10, 2007 for each affected type design, 
and submit those ICA for review and approval by the FAA Oversight 
Office. For purposes of this section, the ``representative airplane'' is 
the configuration of each model series airplane that incorporates all 
variations of EWIS used in production on that series airplane, and all 
TC-holder-designed modifications mandated by airworthiness directive as 
of the effective date of this rule. Each person specified in paragraph 
(d) of this section must also review any fuel tank system ICA developed 
by that person to comply with SFAR 88 to ensure compatibility with the 
EWIS ICA, including minimizing redundant requirements.
    (c) Applicants for amendments to type certificates and supplemental 
type certificates, as identified in paragraph (d) of this section, must:
    (1) Evaluate whether the design change for which approval is sought 
necessitates a revision to the ICA required by paragraph (b) of this 
section to comply with the requirements of Appendix H, paragraphs 
H25.5(a)(1) and (b). If so, the applicant must develop and submit the 
necessary revisions for review and approval by the FAA Oversight Office.
    (2) Ensure that any revised EWIS ICA remain compatible with any fuel 
tank system ICA previously developed to

[[Page 644]]

comply with SFAR 88 and any redundant requirements between them are 
minimized.
    (d) The following persons must comply with the requirements of 
paragraph (b) or (c) of this section, as applicable, before the dates 
specified.
    (1) Holders of type certificates (TC): December 10, 2009.
    (2) Applicants for TCs, and amendments to TCs (including service 
bulletins describing design changes), if the date of application was 
before December 10, 2007 and the certificate was issued on or after 
December 10, 2007: December 10, 2009 or the date the certificate is 
issued, whichever occurs later.
    (3) Unless compliance with Sec. 25.1729 of this subchapter is 
required or elected, applicants for amendments to TCs, if the 
application was filed on or after December 10, 2007: December 10, 2009, 
or the date of approval of the certificate, whichever occurs later.
    (4) Applicants for supplemental type certificates (STC), including 
changes to existing STCs, if the date of application was before December 
10, 2007 and the certificate was issued on or after December 10, 2007: 
June 7, 2010, or the date of approval of the certificate, whichever 
occurs later.
    (5) Unless compliance with Sec. 25.1729 of this subchapter is 
required or elected, applicants for STCs, including changes to existing 
STCs, if the application was filed on or after December 10, 2007, June 
7, 2010, or the date of approval of the certificate, whichever occurs 
later.
    (e) Each person identified in paragraphs (d)(1), (d)(2), and (d)(4) 
of this section must submit to the FAA Oversight Office for approval a 
compliance plan by March 10, 2008. The compliance plan must include the 
following information:
    (1) A proposed project schedule, identifying all major milestones, 
for meeting the compliance dates specified in paragraph (d) of this 
section.
    (2) A proposed means of compliance with this section, identifying 
all required submissions, including all compliance items as mandated in 
part 25, Appendix H paragraphs H25.5(a)(1) and (b) of this subchapter in 
effect on December 10, 2007, and all data to be developed to 
substantiate compliance.
    (3) A proposal for submitting a draft of all compliance items 
required by paragraph (e)(2) of this section for review by the FAA 
Oversight Office not less than 60 days before the compliance time 
specified in paragraph (d) of this section.
    (4) A proposal for how the approved ICA will be made available to 
affected persons.
    (f) Each person specified in paragraph (e) must implement the 
compliance plan, or later approved revisions, as approved in compliance 
with paragraph (e) of this section.
    (g) This section does not apply to the following airplane models:

(1) Lockheed L-188
(2) Bombardier CL-44
(3) Mitsubishi YS-11
(4) British Aerospace BAC 1-11
(5) Concorde
(6) deHavilland D.H. 106 Comet 4C
(7) VFW--Vereinigte Flugtechnische Werk VFW-614
(8) Illyushin Aviation IL 96T
(9) Bristol Aircraft Britannia 305
(10) Handley Page Herald Type 300
(11) Avions Marcel Dassault--Breguet Aviation Mercure 100C
(12) Airbus Caravelle
(13) Lockheed L-300

[Amdt. 26-0, 72 FR 63409, Nov. 8, 2007; 72 FR 68618, Dec. 5, 2007]



        Subpart C_Aging Airplane Safety_Widespread Fatigue Damage

    Source: Docket No. FAA-2006-24281, 75 FR 69782, Nov. 15, 2010, 
unless otherwise noted.



Sec. 26.21  Limit of validity.

    (a) Applicability. Except as provided in paragraph (g) of this 
section, this section applies to transport category, turbine-powered 
airplanes with a maximum takeoff gross weight greater than 75,000 pounds 
and a type certificate issued after January 1, 1958, regardless of 
whether the maximum takeoff gross weight is a result of an original type 
certificate or a later design change. This section also applies to 
transport category, turbine-powered airplanes with a type certificate 
issued after January 1, 1958, if a design change approval for which 
application is made

[[Page 645]]

after January 14, 2011 has the effect of reducing the maximum takeoff 
gross weight from greater than 75,000 pounds to 75,000 pounds or less.
    (b) Limit of validity. Each person identified in paragraph (c) of 
this section must comply with the following requirements:
    (1) Establish a limit of validity of the engineering data that 
supports the structural maintenance program (hereafter referred to as 
LOV) that corresponds to the period of time, stated as a number of total 
accumulated flight cycles or flight hours or both, during which it is 
demonstrated that widespread fatigue damage will not occur in the 
airplane. This demonstration must include an evaluation of airplane 
structural configurations and be supported by test evidence and analysis 
at a minimum and, if available, service experience, or service 
experience and teardown inspection results, of high-time airplanes of 
similar structural design, accounting for differences in operating 
conditions and procedures. The airplane structural configurations to be 
evaluated include--
    (i) All model variations and derivatives approved under the type 
certificate; and
    (ii) All structural modifications to and replacements for the 
airplane structural configurations specified in paragraph (b)(1)(i) of 
this section, mandated by airworthiness directives as of January 14, 
2011.
    (2) If the LOV depends on performance of maintenance actions for 
which service information has not been mandated by airworthiness 
directive as of January 14, 2011, submit the following to the FAA 
Oversight Office:
    (i) For those maintenance actions for which service information has 
been issued as of the applicable compliance date specified in paragraph 
(c) of this section, a list identifying each of those actions.
    (ii) For those maintenance actions for which service information has 
not been issued as of the applicable compliance date specified in 
paragraph (c) of this section, a list identifying each of those actions 
and a binding schedule for providing in a timely manner the necessary 
service information for those actions. Once the FAA Oversight Office 
approves this schedule, each person identified in paragraph (c) of this 
section must comply with that schedule.
    (3) Unless previously accomplished, establish an Airworthiness 
Limitations section (ALS) for each airplane structural configuration 
evaluated under paragraph (b)(1) of this section.
    (4) Incorporate the applicable LOV established under paragraph 
(b)(1) of this section into the ALS for each airplane structural 
configuration evaluated under paragraph (b)(1) and submit it to the FAA 
Oversight Office for approval.
    (c) Persons who must comply and compliance dates. The following 
persons must comply with the requirements of paragraph (b) of this 
section by the specified date.
    (1) Holders of type certificates (TC) of airplane models identified 
in Table 1 of this section: No later than the applicable date identified 
in Table 1 of this section.
    (2) Applicants for TCs, if the date of application was before 
January 14, 2011: No later than the latest of the following dates:
    (i) January 14, 2016;
    (ii) The date the certificate is issued; or
    (iii) The date specified in the plan approved under Sec. 25.571(b) 
for completion of the full-scale fatigue testing and demonstrating that 
widespread fatigue damage will not occur in the airplane structure.
    (3) Applicants for amendments to TCs, with the exception of 
amendments to TCs specified in paragraphs (c)(6) or (c)(7) of this 
section, if the original TC was issued before January 14, 2011: No later 
than the latest of the following dates:
    (i) January 14, 2016;
    (ii) The date the amended certificate is issued; or
    (iii) The date specified in the plan approved under Sec. 25.571(b) 
for completion of the full-scale fatigue testing and demonstrating that 
widespread fatigue damage will not occur in the airplane structure.
    (4) Applicants for amendments to TCs, with the exception of 
amendments to TCs specified in paragraphs (c)(6) or (c)(7) of this 
section, if the application for the original TC was made before

[[Page 646]]

January 14, 2011 but the TC was not issued before January 14, 2011: No 
later than the latest of the following dates:
    (i) January 14, 2016;
    (ii) The date the amended certificate is issued; or
    (iii) The date specified in the plan approved under Sec. 25.571(b) 
for completion of the full-scale fatigue testing and demonstrating that 
widespread fatigue damage will not occur in the airplane structure.
    (5) Holders of either supplemental type certificates (STCs) or 
amendments to TCs that increase maximum takeoff gross weights from 
75,000 pounds or less to greater than 75,000 pounds: No later than July 
14, 2012.
    (6) Applicants for either STCs or amendments to TCs that increase 
maximum takeoff gross weights from 75,000 pounds or less to greater than 
75,000 pounds: No later than the latest of the following dates:
    (i) July 14, 2012;
    (ii) The date the certificate is issued; or
    (iii) The date specified in the plan approved under Sec. 25.571(b) 
for completion of the full-scale fatigue testing and demonstrating that 
widespread fatigue damage will not occur in the airplane structure.
    (7) Applicants for either STCs or amendments to TCs that decrease 
maximum takeoff gross weights from greater than 75,000 pounds to 75,000 
pounds or less, if the date of application was after January 14, 2011: 
No later than the latest of the following dates:
    (i) July 14, 2012;
    (ii) The date the certificate is issued; or
    (iii) The date specified in the plan approved under Sec. 25.571(b) 
for completion of the full-scale fatigue testing and demonstrating that 
widespread fatigue damage will not occur in the airplane structure.
    (d) Compliance plan. Each person identified in paragraph (e) of this 
section must submit a compliance plan consisting of the following:
    (1) A proposed project schedule, identifying all major milestones, 
for meeting the compliance dates specified in paragraph (c) of this 
section.
    (2) A proposed means of compliance with paragraphs (b)(1) through 
(b)(4) of this section.
    (3) A proposal for submitting a draft of all compliance items 
required by paragraph (b) of this section for review by the FAA 
Oversight Office not less than 60 days before the compliance date 
specified in paragraph (c) of this section, as applicable.
    (4) A proposal for how the LOV will be distributed.
    (e) Compliance dates for compliance plans. The following persons 
must submit the compliance plan described in paragraph (d) of this 
section to the FAA Oversight Office by the specified date.
    (1) Holders of type certificates: No later than April 14, 2011.
    (2) Applicants for TCs and amendments to TCs, with the exception of 
amendments to TCs specified in paragraphs (e)(4), (e)(5), or (e)(6) of 
this section, if the date of application was before January 14, 2011 but 
the TC or TC amendment was not issued before January 14, 2011: No later 
than April 14, 2011.
    (3) Holders of either supplemental type certificates or amendments 
to TCs that increase maximum takeoff gross weights from 75,000 pounds or 
less to greater than 75,000 pounds: No later than April 14, 2011.
    (4) Applicants for either STCs or amendments to TCs that increase 
maximum takeoff gross weights from 75,000 pounds or less to greater than 
75,000 pounds, if the date of application was before January 14, 2011: 
No later than April 14, 2011.
    (5) Applicants for either STCs or amendments to TCs that increase 
maximum takeoff gross weights from 75,000 pounds or less to greater than 
75,000 pounds, if the date of application is on or after January 14, 
2011: Within 90 days after the date of application.
    (6) Applicants for either STCs or amendments to TCs that decrease 
maximum takeoff gross weights from greater than 75,000 pounds to 75,000 
pounds or less, if the date of application is on or after January 14, 
2011: Within 90 days after the date of application.
    (f) Compliance plan implementation. Each affected person must 
implement

[[Page 647]]

the compliance plan as approved in compliance with paragraph (d) of this 
section.
    (g) Exceptions. This section does not apply to the following 
airplane models:
    (1) Bombardier BD-700.
    (2) Bombardier CL-44.
    (3) Gulfstream GV.
    (4) Gulfstream GV-SP.
    (5) British Aerospace, Aircraft Group, and Societe Nationale 
Industrielle Aerospatiale Concorde Type 1.
    (6) British Aerospace (Commercial Aircraft) Ltd., Armstrong 
Whitworth Argosy A.W. 650 Series 101.
    (7) British Aerospace Airbus, Ltd., BAC 1-11.
    (8) BAE Systems (Operations) Ltd., BAe 146.
    (9) BAE Systems (Operations) Ltd., Avro 146.
    (10) Lockheed 300-50A01 (USAF C141A).
    (11) Boeing 707.
    (12) Boeing 720.
    (13) deHavilland D.H. 106 Comet 4C.
    (14) Ilyushin Aviation IL-96T.
    (15) Bristol Aircraft Britannia 305.
    (16) Avions Marcel Dassault-Breguet Aviation Mercure 100C.
    (17) Airbus Caravelle.
    (18) D & R Nevada, LLC, Convair Model 22.
    (19) D & R Nevada, LLC, Convair Model 23M.

            Table 1--Compliance Dates for Affected Airplanes
------------------------------------------------------------------------
                                                      Compliance date--
     Airplane model (all existing \1\ models)           (months after
                                                      January 14, 2011)
------------------------------------------------------------------------
Airbus:
    A300 Series...................................                    18
    A310 Series, A300-600 Series..................                    48
    A318 Series...................................                    48
    A319 Series...................................                    48
    A320 Series...................................                    48
    A321 Series...................................                    48
    A330-200, -200 Freighter, -300 Series.........                    48
    A340-200, -300, -500, -600 Series.............                    48
    A380-800 Series...............................                    60
Boeing:
    717...........................................                    48
    727 (all series)..............................                    18
    737 (Classics): 737-100, -200, -200C, -300, -                     18
     400, -500....................................
    737 (NG): 737-600, -700, -700C, -800, -900, -                     48
     900ER........................................
    747 (Classics): 747-100, -100B, -100B SUD, -                      18
     200B, -200C, -200F, -300, 747SP, 747SR.......
    747-400: 747-400, -400D, -400F................                    48
    757...........................................                    48
    767...........................................                    48
    777-200, -300.................................                    48
    777-200LR, 777-300ER, 777F....................                    60
Bombardier:
    CL-600: 2D15 (Regional Jet Series 705), 2D24                      60
     (Regional Jet Series 900)....................
Embraer:
    ERJ 170.......................................                    60
    ERJ 190.......................................                    60
Fokker:
    F.28 Mark 0070, Mark 0100.....................                    18
Lockheed:
    L-1011........................................                    18
    188...........................................                    18
    382 (all series)..............................                    18
McDonnell Douglas:
    DC-8, -8F.....................................                    18
    DC-9..........................................                    18
    MD-80 (DC-9-81, -82, -83, -87, MD-88).........                    18
    MD-90.........................................                    48
    DC-10.........................................                    18
    MD-10.........................................                    48
    MD-11, -11F...................................                    48
    All Other Airplane Models Listed on a Type                        60
     Certificate as of January 14, 2011...........
------------------------------------------------------------------------
\1\ Type certificated as of January 14, 2011.


[[Page 648]]


[Doc. No. FAA-2006-24281, 75 FR 69782, Nov. 15, 2010, as amended at 77 
FR 30878, May 24, 2012]



Sec. 26.23  Extended limit of validity.

    (a) Applicability. Any person may apply to extend a limit of 
validity of the engineering data that supports the structural 
maintenance program (hereafter referred to as LOV) approved under Sec. 
25.571 of this subchapter, Sec. 26.21, or this section. Extending an 
LOV is a major design change. The applicant must comply with the 
relevant provisions of subparts D or E of part 21 of this subchapter and 
paragraph (b) of this section.
    (b) Extended limit of validity. Each person applying for an extended 
LOV must comply with the following requirements:
    (1) Establish an extended LOV that corresponds to the period of 
time, stated as a number of total accumulated flight cycles or flight 
hours or both, during which it is demonstrated that widespread fatigue 
damage will not occur in the airplane. This demonstration must include 
an evaluation of airplane structural configurations and be supported by 
test evidence and analysis at a minimum and, if available, service 
experience, or service experience and teardown inspection results, of 
high-time airplanes of similar structural design, accounting for 
differences in operating conditions and procedures. The airplane 
structural configurations to be evaluated include--
    (i) All model variations and derivatives approved under the type 
certificate for which approval for an extension is sought; and
    (ii) All structural modifications to and replacements for the 
airplane structural configurations specified in paragraph (b)(1)(i) of 
this section, mandated by airworthiness directive, up to the date of 
approval of the extended LOV.
    (2) Establish a revision or supplement, as applicable, to the 
Airworthiness Limitations section (ALS) of the Instructions for 
Continued Airworthiness required by Sec. 25.1529 of this subchapter, 
and submit it to the FAA Oversight Office for approval. The revised ALS 
or supplement to the ALS must include the applicable extended LOV 
established under paragraph (b)(1) of this section.
    (3) Develop the maintenance actions determined by the WFD evaluation 
performed in paragraph (b)(1) of this section to be necessary to 
preclude WFD from occurring before the airplane reaches the proposed 
extended LOV. These maintenance actions must be documented as 
airworthiness limitation items in the ALS and submitted to the FAA 
Oversight Office for approval.



                    Subpart D_Fuel Tank Flammability

    Source: Docket No. FAA-2005-22997, 73 FR 42499, July 21, 2008, 
unless otherwise noted.



Sec. 26.31  Definitions.

    For purposes of this subpart--
    (a) Fleet Average Flammability Exposure has the meaning defined in 
Appendix N of part 25 of this chapter.
    (b) Normally Emptied means a fuel tank other than a Main Fuel Tank. 
Main Fuel Tank is defined in 14 CFR 25.981(b).



Sec. 26.33  Holders of type certificates: Fuel tank flammability.

    (a) Applicability. This section applies to U.S. type certificated 
transport category, turbine-powered airplanes, other than those designed 
solely for all-cargo operations, for which the State of Manufacture 
issued the original certificate of airworthiness or export airworthiness 
approval on or after January 1, 1992, that, as a result of original type 
certification or later increase in capacity have:
    (1) A maximum type-certificated passenger capacity of 30 or more, or
    (2) A maximum payload capacity of 7,500 pounds or more.
    (b) Flammability Exposure Analysis. (1) General. Within 150 days 
after December 26, 2008, holders of type certificates must submit for 
approval to the FAA Oversight Office a flammability exposure analysis of 
all fuel tanks defined in the type design, as well as all design 
variations approved under the type certificate that affect flammability 
exposure. This analysis must be conducted in accordance with Appendix N 
of part 25 of this chapter.

[[Page 649]]

    (2) Exception. This paragraph (b) does not apply to--
    (i) Fuel tanks for which the type certificate holder has notified 
the FAA under paragraph (g) of this section that it will provide design 
changes and service instructions for Flammability Reduction Means or an 
Ignition Mitigation Means (IMM) meeting the requirements of paragraph 
(c) of this section.
    (ii) Fuel tanks substantiated to be conventional unheated aluminum 
wing tanks.
    (c) Design Changes. For fuel tanks with a Fleet Average Flammability 
Exposure exceeding 7 percent, one of the following design changes must 
be made.
    (1) Flammability Reduction Means (FRM). A means must be provided to 
reduce the fuel tank flammability.
    (i) Fuel tanks that are designed to be Normally Emptied must meet 
the flammability exposure criteria of Appendix M of part 25 of this 
chapter if any portion of the tank is located within the fuselage 
contour.
    (ii) For all other fuel tanks, the FRM must meet all of the 
requirements of Appendix M of part 25 of this chapter, except, instead 
of complying with paragraph M25.1 of this appendix, the Fleet Average 
Flammability Exposure may not exceed 7 percent.
    (2) Ignition Mitigation Means (IMM). A means must be provided to 
mitigate the effects of an ignition of fuel vapors within the fuel tank 
such that no damage caused by an ignition will prevent continued safe 
flight and landing.
    (d) Service Instructions. No later than December 27, 2010, holders 
of type certificates required by paragraph (c) of this section to make 
design changes must meet the requirements specified in either paragraph 
(d)(1) or (d)(2) of this section. The required service instructions must 
identify each airplane subject to the applicability provisions of 
paragraph (a) of this section.
    (1) FRM. The type certificate holder must submit for approval by the 
FAA Oversight Office design changes and service instructions for 
installation of fuel tank flammability reduction means (FRM) meeting the 
criteria of paragraph (c) of this section.
    (2) IMM. The type certificate holder must submit for approval by the 
FAA Oversight Office design changes and service instructions for 
installation of fuel tank IMM that comply with 14 CFR 25.981(c) in 
effect on December 26, 2008.
    (e) Instructions for Continued Airworthiness (ICA). No later than 
December 27, 2010, holders of type certificates required by paragraph 
(c) of this section to make design changes must submit for approval by 
the FAA Oversight Office, critical design configuration control 
limitations (CDCCL), inspections, or other procedures to prevent 
increasing the flammability exposure of any tanks equipped with FRM 
above that permitted under paragraph (c)(1) of this section and to 
prevent degradation of the performance of any IMM provided under 
paragraph (c)(2) of this section. These CDCCL, inspections, and 
procedures must be included in the Airworthiness Limitations Section 
(ALS) of the ICA required by 14 CFR 25.1529 or paragraph (f) of this 
section. Unless shown to be impracticable, visible means to identify 
critical features of the design must be placed in areas of the airplane 
where foreseeable maintenance actions, repairs, or alterations may 
compromise the critical design configuration limitations. These visible 
means must also be identified as a CDCCL.
    (f) Airworthiness Limitations. Unless previously accomplished, no 
later than December 27, 2010, holders of type certificates affected by 
this section must establish an ALS of the maintenance manual or ICA for 
each airplane configuration evaluated under paragraph (b)(1) of this 
section and submit it to the FAA Oversight Office for approval. The ALS 
must include a section that contains the CDCCL, inspections, or other 
procedures developed under paragraph (e) of this section.
    (g) Compliance Plan for Flammability Exposure Analysis. Within 90 
days after December 26, 2008, each holder of a type certificate required 
to comply with paragraph (b) of this section must submit to the FAA 
Oversight Office a compliance plan consisting of the following:
    (1) A proposed project schedule for submitting the required 
analysis, or a determination that compliance with

[[Page 650]]

paragraph (b) of this section is not required because design changes and 
service instructions for FRM or IMM will be developed and made available 
as required by this section.
    (2) A proposed means of compliance with paragraph (b) of this 
section, if applicable.
    (h) Compliance Plan for Design Changes and Service Instructions. 
Within 210 days after December 26, 2008, each holder of a type 
certificate required to comply with paragraph (d) of this section must 
submit to the FAA Oversight Office a compliance plan consisting of the 
following:
    (1) A proposed project schedule, identifying all major milestones, 
for meeting the compliance dates specified in paragraphs (d), (e) and 
(f) of this section.
    (2) A proposed means of compliance with paragraphs (d), (e) and (f) 
of this section.
    (3) A proposal for submitting a draft of all compliance items 
required by paragraphs (d), (e) and (f) of this section for review by 
the FAA Oversight Office not less than 60 days before the compliance 
times specified in those paragraphs.
    (4) A proposal for how the approved service information and any 
necessary modification parts will be made available to affected persons.
    (i) Each affected type certificate holder must implement the 
compliance plans, or later revisions, as approved under paragraph (g) 
and (h) of this section.

[Doc. No. FAA-2005-22997, 73 FR 42499, July 21, 2008, as amended by 
Amdt. 26-3, 74 FR 31619, July 2, 2009]



Sec. 26.35  Changes to type certificates affecting fuel tank flammability.

    (a) Applicability. This section applies to holders and applicants 
for approvals of the following design changes to any airplane subject to 
14 CFR 26.33(a):
    (1) Any fuel tank designed to be Normally Emptied if the fuel tank 
installation was approved pursuant to a supplemental type certificate or 
a field approval before December 26, 2008;
    (2) Any fuel tank designed to be Normally Emptied if an application 
for a supplemental type certificate or an amendment to a type 
certificate was made before December 26, 2008 and if the approval was 
not issued before December 26, 2008; and
    (3) If an application for a supplemental type certificate or an 
amendment to a type certificate is made on or after December 26, 2008, 
any of the following design changes:
    (i) Installation of a fuel tank designed to be Normally Emptied,
    (ii) Changes to existing fuel tank capacity, or
    (iii) Changes that may increase the flammability exposure of an 
existing fuel tank for which FRM or IMM is required by Sec. 26.33(c).
    (b) Flammability Exposure Analysis--(1) General. By the times 
specified in paragraphs (b)(1)(i) and (b)(1)(ii) of this section, each 
person subject to this section must submit for approval a flammability 
exposure analysis of the auxiliary fuel tanks or other affected fuel 
tanks, as defined in the type design, to the FAA Oversight Office. This 
analysis must be conducted in accordance with Appendix N of part 25 of 
this chapter.
    (i) Holders of supplemental type certificates and field approvals: 
Within 12 months of December 26, 2008,
    (ii) Applicants for supplemental type certificates and for 
amendments to type certificates: Within 12 months after December 26, 
2008, or before the certificate is issued, whichever occurs later.
    (2) Exception. This paragraph does not apply to--
    (i) Fuel tanks for which the type certificate holder, supplemental 
type certificate holder, or field approval holder has notified the FAA 
under paragraph (f) of this section that it will provide design changes 
and service instructions for an IMM meeting the requirements of Sec. 
25.981(c) in effect December 26, 2008; and
    (ii) Fuel tanks substantiated to be conventional unheated aluminum 
wing tanks.
    (c) Impact Assessment. By the times specified in paragraphs (c)(1) 
and (c)(2) of this section, each person subject to paragraph (a)(1) of 
this section holding an approval for installation of a Normally Emptied 
fuel tank on an airplane model listed in Table 1 of this section, and 
each person subject to paragraph

[[Page 651]]

(a)(3)(iii) of this section, must submit for approval to the FAA 
Oversight Office an assessment of the fuel tank system, as modified by 
their design change. The assessment must identify any features of the 
design change that compromise any critical design configuration control 
limitation (CDCCL) applicable to any airplane on which the design change 
is eligible for installation.
    (1) Holders of supplemental type certificates and field approvals: 
Before June 26, 2011.
    (2) Applicants for supplemental type certificates and for amendments 
to type certificates: Before June 26, 2011 or before the certificate is 
issued, whichever occurs later.

                                 Table 1
------------------------------------------------------------------------
 
-------------------------------------------------------------------------
                              Model--Boeing
------------------------------------------------------------------------
747 Series
737 Series
777 Series
767 Series
757 Series
------------------------------------------------------------------------
                              Model--Airbus
------------------------------------------------------------------------
A318, A319, A320, A321 Series
A300, A310 Series
A330, A340 Series
------------------------------------------------------------------------

    (d) Design Changes and Service Instructions. By the times specified 
in paragraph (e) of this section, each person subject to this section 
must meet the requirements of paragraphs (d)(1) or (d)(2) of this 
section, as applicable.
    (1) For holders and applicants subject to paragraph (a)(1) or 
(a)(3)(iii) of this section, if the assessment required by paragraph (c) 
of this section identifies any features of the design change that 
compromise any CDCCL applicable to any airplane on which the design 
change is eligible for installation, the holder or applicant must submit 
for approval by the FAA Oversight Office design changes and service 
instructions for Flammability Impact Mitigation Means (FIMM) that would 
bring the design change into compliance with the CDCCL. Any fuel tank 
modified as required by this paragraph must also be evaluated as 
required by paragraph (b) of this section.
    (2) Applicants subject to paragraph (a)(2), or (a)(3)(i) of this 
section must comply with the requirements of 14 CFR 25.981, in effect on 
December 26, 2008.
    (3) Applicants subject to paragraph (a)(3)(ii) of this section must 
comply with the requirements of 14 CFR 26.33.
    (e) Compliance Times for Design Changes and Service Instructions. 
The following persons subject to this section must comply with the 
requirements of paragraph (d) of this section at the specified times.
    (1) Holders of supplemental type certificates and field approvals: 
Before December 26, 2012.
    (2) Applicants for supplemental type certificates and for amendments 
to type certificates: Before December 26, 2012, or before the 
certificate is issued, whichever occurs later.
    (f) Compliance Planning. By the applicable date specified in Table 2 
of this section, each person subject to paragraph (a)(1) of this section 
must submit for approval by the FAA Oversight Office compliance plans 
for the flammability exposure analysis required by paragraph (b) of this 
section, the impact assessment required by paragraph (c) of this 
section, and the design changes and service instructions required by 
paragraph (d) of this section. Each person's compliance plans must 
include the following:
    (1) A proposed project schedule for submitting the required analysis 
or impact assessment.
    (2) A proposed means of compliance with paragraph (d) of this 
section.
    (3) For the requirements of paragraph (d) of this section, a 
proposal for submitting a draft of all design changes, if any are 
required, and Airworthiness Limitations (including CDCCLs) for review by 
the FAA Oversight Office not less than 60 days before the compliance 
time specified in paragraph (e) of this section.
    (4) For the requirements of paragraph (d) of this section, a 
proposal for how the approved service information and any necessary 
modification parts will be made available to affected persons.

[[Page 652]]



                                       Table 2--Compliance Planning Dates
----------------------------------------------------------------------------------------------------------------
                                    Flammability exposure    Impact assessment      Design changes and service
                                        analysis plan              plan                 instructions plan
----------------------------------------------------------------------------------------------------------------
STC and Field Approval Holders....  March 26, 2009.......  February 26, 2011...  August 26, 2011.
----------------------------------------------------------------------------------------------------------------

    (g) Each person subject to this section must implement the 
compliance plans, or later revisions, as approved under paragraph (f) of 
this section.

[Doc. No. FAA-2005-22997, 73 FR 42499, July 21, 2008, as amended by 
Amdt. 26-3, 74 FR 31619, July 2, 2009]



Sec. 26.37  Pending type certification projects: Fuel tank flammability.

    (a) Applicability. This section applies to any new type certificate 
for a transport category airplane, if the application was made before 
December 26, 2008, and if the certificate was not issued before December 
26, 2008. This section applies only if the airplane would have--
    (1) A maximum type-certificated passenger capacity of 30 or more, or
    (2) A maximum payload capacity of 7,500 pounds or more.
    (b) If the application was made on or after June 6, 2001, the 
requirements of 14 CFR 25.981 in effect on December 26, 2008, apply.

[Doc. No. FAA-2005-22997, 73 FR 42499, July 21, 2008, as amended by 
Amdt. 26-3, 74 FR 31619, July 2, 2009]



Sec. 26.39  Newly produced airplanes: Fuel tank flammability.

    (a) Applicability: This section applies to Boeing model airplanes 
specified in Table 1 of this section, including passenger and cargo 
versions of each model, when application is made for original 
certificates of airworthiness or export airworthiness approvals after 
December 27, 2010.

                                 Table 1
------------------------------------------------------------------------
                              Model--Boeing
-------------------------------------------------------------------------
747 Series
737 Series
777 Series
767 Series
------------------------------------------------------------------------

    (b) Any fuel tank meeting all of the criteria stated in paragraphs 
(b)(1), (b)(2) and (b)(3) of this section must have flammability 
reduction means (FRM) or ignition mitigation means (IMM) that meet the 
requirements of 14 CFR 25.981 in effect on December 26, 2008.
    (1) The fuel tank is Normally Emptied.
    (2) Any portion of the fuel tank is located within the fuselage 
contour.
    (3) The fuel tank exceeds a Fleet Average Flammability Exposure of 7 
percent.
    (c) All other fuel tanks that exceed an Fleet Average Flammability 
Exposure of 7 percent must have an IMM that meets 14 CFR 25.981(d) in 
effect on December 26, 2008, or an FRM that meets all of the 
requirements of Appendix M to this part, except instead of complying 
with paragraph M25.1 of that appendix, the Fleet Average Flammability 
Exposure may not exceed 7 percent.

[Doc. No. FAA-2005-22997, 73 FR 42499, July 21, 2008, as amended by 
Amdt. 26-3, 74 FR 31619, July 2, 2009]



 Subpart E_Aging Airplane Safety_Damage Tolerance Data for Repairs and 
                               Alterations

    Source: Docket No. FAA-2005-21693, 72 FR 70505, Dec. 12, 2007, 
unless otherwise noted.



Sec. 26.41  Definitions.

    Affects (or Affected) means structure has been physically repaired, 
altered, or modified, or the structural loads acting on the structure 
have been increased or redistributed.
    Baseline structure means structure that is designed under the 
original type certificate or amended type certificate for that airplane 
model.
    Damage Tolerance Evaluation (DTE) means a process that leads to a 
determination of maintenance actions necessary to detect or preclude 
fatigue cracking that could contribute to a

[[Page 653]]

catastrophic failure. As applied to repairs and alterations, a DTE 
includes the evaluation both of the repair or alteration and of the 
fatigue critical structure affected by the repair or alteration.
    Damage Tolerance Inspection (DTI) means the inspection developed as 
a result of a DTE. A DTI includes the areas to be inspected, the 
inspection method, the inspection procedures, including acceptance and 
rejection criteria, the threshold, and any repeat intervals associated 
with those inspections. The DTI may specify a time limit when a repair 
or alteration needs to be replaced or modified. If the DTE concludes 
that DT-based supplemental structural inspections are not necessary, the 
DTI contains a statement to that effect.
    DT data mean DTE documentation and the DTI.
    DTE documentation means data that identify the evaluated fatigue 
critical structure, the basic assumptions applied in a DTE, and the 
results of a DTE.
    Fatigue critical structure means airplane structure that is 
susceptible to fatigue cracking that could contribute to a catastrophic 
failure, as determined in accordance with Sec. 25.571 of this chapter. 
Fatigue critical structure includes structure, which, if repaired or 
altered, could be susceptible to fatigue cracking and contribute to a 
catastrophic failure. Such structure may be part of the baseline 
structure or part of an alteration.
    Implementation schedule consists of documentation that establishes 
the timing for accomplishing the necessary actions for developing DT 
data for repairs and alterations, and for incorporating those data into 
an operator's continuing airworthiness maintenance program. The 
documentation must identify times when actions must be taken as specific 
numbers of airplane flight hours, flight cycles, or both.
    Published repair data mean instructions for accomplishing repairs, 
which are published for general use in structural repair manuals and 
service bulletins (or equivalent types of documents).



Sec. 26.43  Holders of and applicants for type certificates--Repairs.

    (a) Applicability. Except as specified in paragraph (g) of this 
section, this section applies to transport category, turbine powered 
airplane models with a type certificate issued after January 1, 1958, 
that as a result of original type certification or later increase in 
capacity have--
    (1) A maximum type certificated passenger seating capacity of 30 or 
more; or
    (2) A maximum payload capacity of 7,500 pounds or more.
    (b) List of fatigue critical baseline structure. For airplanes 
specified in paragraph (a) of this section, the holder of or applicant 
for a type certificate must--
    (1) Identify fatigue critical baseline structure for all airplane 
model variations and derivatives approved under the type certificate; 
and
    (2) Develop and submit to the FAA Oversight Office for review and 
approval, a list of the structure identified under paragraph (b)(1) of 
this section and, upon approval, make the list available to persons 
required to comply with Sec. 26.47 and Sec. Sec. 121.1109 and 129.109 
of this chapter.
    (c) Existing and future published repair data. For repair data 
published by a holder of a type certificate that is current as of 
January 11, 2008 and for all later published repair data, the holder of 
a type certificate must--
    (1) Review the repair data and identify each repair specified in the 
data that affects fatigue critical baseline structure identified under 
paragraph (b)(1) of this section;
    (2) Perform a DTE and develop the DTI for each repair identified 
under paragraph (c)(1) of this section, unless previously accomplished;
    (3) Submit the DT data to the FAA Oversight Office or its properly 
authorized designees for review and approval; and
    (4) Upon approval, make the DTI available to persons required to 
comply with Sec. Sec. 121.1109 and 129.109 of this chapter.
    (d) Future repair data not published. For repair data developed by a 
holder of a type certificate that are approved

[[Page 654]]

after January 11, 2008 and are not published, the type certificate 
holder must accomplish the following for repairs specified in the repair 
data that affect fatigue critical baseline structure:
    (1) Perform a DTE and develop the DTI.
    (2) Submit the DT data required in paragraph (d)(1) of this section 
for review and approval by the FAA Oversight Office or its properly 
authorized designees.
    (3) Upon approval, make the approved DTI available to persons 
required to comply with Sec. Sec. 121.1109 and 129.109 of this chapter.
    (e) Repair evaluation guidelines. Except for airplane models whose 
type certificate is issued after January 11, 2008, holders of a type 
certificate for each airplane model subject to this section must--
    (1) Develop repair evaluation guidelines for operators' use that 
include--
    (i) A process for conducting surveys of affected airplanes that will 
enable identification and documentation of all existing repairs that 
affect fatigue critical baseline structure identified under paragraph 
(b)(1) of this section and Sec. 26.45(b)(2);
    (ii) A process that will enable operators to obtain the DTI for 
repairs identified under paragraph (e)(1)(i) of this section; and
    (iii) An implementation schedule for repairs covered by the repair 
evaluation guidelines. The implementation schedule must identify times 
when actions must be taken as specific numbers of airplane flight hours, 
flight cycles, or both.
    (2) Submit the repair evaluation guidelines to the FAA Oversight 
Office for review and approval.
    (3) Upon approval, make the guidelines available to persons required 
to comply with Sec. Sec. 121.1109 and 129.109 of this chapter.
    (4) If the guidelines direct the operator to obtain assistance from 
the holder of a type certificate, make such assistance available in 
accordance with the implementation schedule.
    (f) Compliance times. Holders of type certificates must submit the 
following to the FAA Oversight Office or its properly authorized 
designees for review and approval by the specified compliance time:
    (1) The identified list of fatigue critical baseline structure 
required by paragraph (b)(2) of this section must be submitted no later 
than 180 days after January 11, 2008 or before issuance of the type 
certificate, whichever occurs later.
    (2) For published repair data that are current as of January 11, 
2008, the DT data required by paragraph (c)(3) of this section must be 
submitted by June 30, 2009.
    (3) For repair data published after January 11, 2008, the DT data 
required by paragraph (c)(3) of this section must be submitted before 
FAA approval of the repair data.
    (4) For unpublished repair data developed after January 11, 2008, 
the DT data required by paragraph (d)(1) of this section must be 
submitted within 12 months of the airplane's return to service or in 
accordance with a schedule approved by the FAA Oversight Office.
    (5) The repair evaluation guidelines required by paragraph (e)(1) of 
this section must be submitted by December 30, 2009.
    (g) Exceptions. The requirements of this section do not apply to the 
following transport category airplane models:
    (1) Convair CV-240, 340, 440, if modified to include turbine 
engines.
    (2) Vickers Armstrong Viscount, TCDS No. A-814.
    (3) Douglas DC-3, if modified to include turbine engines, TCDS No. 
A-618.
    (4) Bombardier CL-44, TCDS No. 1A20.
    (5) Mitsubishi YS-11, TCDS No. A1PC.
    (6) British Aerospace BAC 1-11, TCDS No. A5EU.
    (7) Concorde, TCDS No. A45EU.
    (8) deHavilland D.H. 106 Comet 4C, TCDS No. 7A10.
    (9) deHavilland DHC-7, TCDS No. A20EA.
    (10) VFW-Vereinigte Flugtechnische Werk VFW-614, TCDS No. A39EU.
    (11) Illyushin Aviation IL 96T, TCDS No. A54NM.
    (12) Bristol Aircraft Britannia 305, TCDS No. 7A2.

[[Page 655]]

    (13) Handley Page Herald Type 300, TCDS No. A21N.
    (14) Avions Marcel Dassault--Breguet Aviation Mercure 100C, TCDS No. 
A40EU.
    (15) Airbus Caravelle, TCDS No. 7A6.
    (16) Lockheed L-300, TCDS No. A2S0.
    (17) Boeing 707-100/-200, TCDS No. 4A21.
    (18) Boeing 707-300/-400, TCDS No. 4A26.
    (19) Boeing 720, TCDS No. 4A28.

[Doc. No. FAA-2005-21693, 72 FR 70505, Dec. 12, 2007, as amended by 
Amdt. 26-4, 75 FR 11734, Mar. 12, 2010]



Sec. 26.45  Holders of type certificates--Alterations and repairs
to alterations.

    (a) Applicability. This section applies to transport category 
airplanes subject to Sec. 26.43.
    (b) Fatigue critical alteration structure. For existing and future 
alteration data developed by the holder of a type certificate, the 
holder must--
    (1) Review alteration data and identify all alterations that affect 
fatigue critical baseline structure identified under Sec. 26.43(b)(1);
    (2) For each alteration identified under paragraph (b)(1) of this 
section, identify any fatigue critical alteration structure;
    (3) Develop and submit to the FAA Oversight Office for review and 
approval a list of the structure identified under paragraph (b)(2) of 
this section; and
    (4) Upon approval, make the list required in paragraph (b)(3) of 
this section available to persons required to comply with Sec. Sec. 
121.1109 and 129.109 of this chapter.
    (c) DT Data. For existing and future alteration data developed by 
the holder of a type certificate that affect fatigue critical baseline 
structure identified under Sec. 26.43(b)(1), unless previously 
accomplished, the holder must--
    (1) Perform a DTE and develop the DTI for the alteration and fatigue 
critical baseline structure that is affected by the alteration;
    (2) Submit the DT data developed in accordance with paragraphs 
(c)(1) of this section to the FAA Oversight Office or its properly 
authorized designees for review and approval; and
    (3) Upon approval, make the DTI available to persons required to 
comply with Sec. Sec. 121.1109 and 129.109 of this chapter.
    (d) DT Data for Repairs Made to Alterations. For existing and future 
repair data developed by a holder of a type certificate, the type 
certificate holder must--
    (1) Review the repair data, and identify each repair that affects 
any fatigue critical alteration structure identified under paragraph 
(b)(2) of this section;
    (2) For each repair identified under paragraph (d)(1) of this 
section, unless previously accomplished, perform a DTE and develop DTI;
    (3) Submit the DT data developed in accordance with paragraph (d)(2) 
of this section to the FAA Oversight Office or its properly authorized 
designees for review and approval; and
    (4) Upon approval, make the DTI available to persons required to 
comply with Sec. Sec. 121.1109 and 129.109 of this chapter.
    (e) Compliance times. Holders of type certificates must submit the 
following to the FAA Oversight Office or its properly authorized 
designees for review and approval by the specified compliance time:
    (1) The list of fatigue critical alteration structure identified 
under paragraph (b)(3) of this section must be submitted--
    (i) No later than 360 days after January 11, 2008, for alteration 
data approved before January 11, 2008.
    (ii) No later than 30 days after March 12, 2010 or before initial 
approval of the alteration data, whichever occurs later, for alteration 
data approved on or after January 11, 2008.
    (2) For alteration data developed and approved before January 11, 
2008, the DT data required by paragraph (c)(2) of this section must be 
submitted by June 30, 2009.
    (3) For alteration data approved on or after January 11, 2008, DT 
data required by paragraph (c)(2) of this section must be submitted 
before initial approval of the alteration data.
    (4) For repair data developed and approved before January 11, 2008, 
the DT data required by paragraph (d)(2) of

[[Page 656]]

this section must be submitted by June 30, 2009.
    (5) For repair data developed and approved after January 11, 2008, 
the DT data required by paragraph (d)(2) of this section must be 
submitted within 12 months after initial approval of the repair data and 
before making the DT data available to persons required to comply with 
Sec. Sec. 121.1109 and 129.109 of this chapter.

[Doc. No. FAA-2005-21693, 72 FR 70505, Dec. 12, 2007, as amended by 
Amdt. 26-4, 75 FR 11734, Mar. 12, 2010]



Sec. 26.47  Holders of and applicants for a supplemental type 
certificate--Alterations and repairs to alterations.

    (a) Applicability. This section applies to transport category 
airplanes subject to Sec. 26.43.
    (b) Fatigue critical alteration structure. For existing structural 
alteration data approved under a supplemental certificate, the holder of 
the supplemental certificate must--
    (1) Review the alteration data and identify all alterations that 
affect fatigue critical baseline structure identified under Sec. 
26.43(b)(1);
    (2) For each alteration identified under paragraph (b)(1) of this 
section, identify any fatigue critical alteration structure;
    (3) Develop and submit to the FAA Oversight Office for review and 
approval a list of the structure identified under paragraph (b)(2) of 
this section; and
    (4) Upon approval, make the list required in paragraph (b)(3) of 
this section available to persons required to comply with Sec. Sec. 
121.1109 and 129.109 of this chapter.
    (c) DT Data. For existing and future alteration data developed by 
the holder of a supplemental type certificate that affect fatigue 
critical baseline structure identified under Sec. 26.43(b)(1), unless 
previously accomplished, the holder of a supplemental type certificate 
must--
    (1) Perform a DTE and develop the DTI for the alteration and fatigue 
critical baseline structure that is affected by the alteration;
    (2) Submit the DT data developed in accordance with paragraphs 
(c)(1) of this section to the FAA Oversight Office or its properly 
authorized designees for review and approval; and
    (3) Upon approval, make the DTI available to persons required to 
comply with Sec. Sec. 121.1109 and 129.109 of this chapter.
    (d) DT Data for Repairs Made to Alterations. For existing and future 
repair data developed by the holder of a supplemental holder of a 
supplemental type certificate, the holder of a supplemental type 
certificate must--
    (1) Review the repair data, and identify each repair that affects 
any fatigue critical alteration structure identified under paragraph 
(b)(2) of this section;
    (2) For each repair identified under paragraph (d)(1) of this 
section, unless previously accomplished, perform a DTE and develop DTI;
    (3) Submit the DT data developed in accordance with paragraph (d)(2) 
of this section to the FAA Oversight Office or its properly authorized 
designees for review and approval; and
    (4) Upon approval, make the DTI available to persons required to 
comply with Sec. Sec. 121.1109 and 129.109 of this chapter.
    (e) Compliance times. Holders of supplemental type certificates must 
submit the following to the FAA Oversight Office or its properly 
authorized designees for review and approval by the specified compliance 
time:
    (1) The list of fatigue critical alteration structure required by 
paragraph (b)(3) of this section must be submitted no later than 360 
days after January 11, 2008.
    (2) For alteration data developed and approved before January 11, 
2008, the DT data required by paragraph (c)(2) of this section must be 
submitted by June 30, 2009.
    (3) For alteration data developed after January 11, 2008, the DT 
data required by paragraph (c)(2) of this section must be submitted 
before approval of the alteration data and making it available to 
persons required to comply with Sec. Sec. 121.1109 and 129.109 of this 
chapter.
    (4) For repair data developed and approved before January 11, 2008, 
the DT data required by paragraph (d)(2) of this section must be 
submitted by June 30, 2009.

[[Page 657]]

    (5) For repair data developed and approved after January 11, 2008, 
the DT data required by paragraph (d)(2) of this section, must be 
submitted within 12 months after initial approval of the repair data and 
before making the DT data available to persons required to comply with 
Sec. Sec. 121.1109 and 129.109 of this chapter.



Sec. 26.49  Compliance plan.

    (a) Compliance plan. Except for applicants for type certificates and 
supplemental type certificates whose applications are submitted after 
January 11, 2008, each person identified in Sec. Sec. 26.43, 26.45, and 
26.47, must submit a compliance plan consisting of the following:
    (1) A project schedule identifying all major milestones for meeting 
the compliance times specified in Sec. Sec. 26.43(f), 26.45(e), and 
26.47(e), as applicable.
    (2) A proposed means of compliance with Sec. Sec. 26.43, 26.45, and 
26.47, as applicable.
    (3) A plan for submitting a draft of all compliance items required 
by this subpart for review by the FAA Oversight Office not less than 60 
days before the applicable compliance date.
    (b) Compliance dates for compliance plans. The following persons 
must submit the compliance plan described in paragraph (a) of this 
section to the FAA Oversight Office for approval on the following 
schedule:
    (1) For holders of type certificates, no later than 90 days after 
January 11, 2008.
    (2) For holders of supplemental type certificates no later than 180 
days after January 11, 2008.
    (3) For applicants for changes to type certificates whose 
application are submitted before January 11, 2008, no later than 180 
days after January 11, 2008.
    (c) Compliance Plan Implementation. Each affected person must 
implement the compliance plan as approved in compliance with paragraph 
(a) of this section.



PART 27_AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT--
Table of Contents



                            Subpart A_General

Sec.
27.1 Applicability.
27.2 Special retroactive requirements.

                            Subpart B_Flight

                                 General

27.21 Proof of compliance.
27.25 Weight limits.
27.27 Center of gravity limits.
27.29 Empty weight and corresponding center of gravity.
27.31 Removable ballast.
27.33 Main rotor speed and pitch limits.

                               Performance

27.45 General.
27.49 Performance at minimum operating speed.
27.51 Takeoff.
27.65 Climb: all engines operating.
27.67 Climb: one engine inoperative.
27.71 Autorotation performance.
27.75 Landing.
27.87 Height-speed envelope.

                         Flight Characteristics

27.141 General.
27.143 Controllability and maneuverability.
27.151 Flight controls.
27.161 Trim control.
27.171 Stability: general.
27.173 Static longitudinal stability.
27.175 Demonstration of static longitudinal stability.
27.177 Static directional stability.

                Ground and Water Handling Characteristics

27.231 General.
27.235 Taxiing condition.
27.239 Spray characteristics.
27.241 Ground resonance.

                    Miscellaneous Flight Requirements

27.251 Vibration.

                     Subpart C_Strength Requirements

                                 General

27.301 Loads.
27.303 Factor of safety.
27.305 Strength and deformation.
27.307 Proof of structure.
27.309 Design limitations.

                              Flight Loads

27.321 General.

[[Page 658]]

27.337 Limit maneuvering load factor.
27.339 Resultant limit maneuvering loads.
27.341 Gust loads.
27.351 Yawing conditions.
27.361 Engine torque.

                    Control Surface and System Loads

27.391 General.
27.395 Control system.
27.397 Limit pilot forces and torques.
27.399 Dual control system.
27.411 Ground clearance: tail rotor guard.
27.427 Unsymmetrical loads.

                              Ground Loads

27.471 General.
27.473 Ground loading conditions and assumptions.
27.475 Tires and shock absorbers.
27.477 Landing gear arrangement.
27.479 Level landing conditions.
27.481 Tail-down landing conditions.
27.483 One-wheel landing conditions.
27.485 Lateral drift landing conditions.
27.493 Braked roll conditions.
27.497 Ground loading conditions: landing gear with tail wheels.
27.501 Ground loading conditions: landing gear with skids.
27.505 Ski landing conditions.

                               Water Loads

27.521 Float landing conditions.

                       Main Component Requirements

27.547 Main rotor structure.
27.549 Fuselage, landing gear, and rotor pylon structures.

                      Emergency Landing Conditions

27.561 General.
27.562 Emergency landing dynamic conditions.
27.563 Structural ditching provisions.

                           Fatigue Evaluation

27.571 Fatigue evaluation of flight structure.
27.573 Damage tolerance and fatigue evaluation of composite rotorcraft 
          structures.

                    Subpart D_Design and Construction

                                 General

27.601 Design.
27.602 Critical parts.
27.603 Materials.
27.605 Fabrication methods.
27.607 Fasteners.
27.609 Protection of structure.
27.610 Lightning and static electricity protection.
27.611 Inspection provisions.
27.613 Material strength properties and design values.
27.619 Special factors.
27.621 Casting factors.
27.623 Bearing factors.
27.625 Fitting factors.
27.629 Flutter.

                                 Rotors

27.653 Pressure venting and drainage of rotor blades.
27.659 Mass balance.
27.661 Rotor blade clearance.
27.663 Ground resonance prevention means.

                             Control Systems

27.671 General.
27.672 Stability augmentation, automatic, and power-operated systems.
27.673 Primary flight control.
27.674 Interconnected controls.
27.675 Stops.
27.679 Control system locks.
27.681 Limit load static tests.
27.683 Operation tests.
27.685 Control system details.
27.687 Spring devices.
27.691 Autorotation control mechanism.
27.695 Power boost and power-operated control system.

                              Landing Gear

27.723 Shock absorption tests.
27.725 Limit drop test.
27.727 Reserve energy absorption drop test.
27.729 Retracting mechanism.
27.731 Wheels.
27.733 Tires.
27.735 Brakes.
27.737 Skis.

                            Floats and Hulls

27.751 Main float buoyancy.
27.753 Main float design.
27.755 Hulls.

                   Personnel and Cargo Accommodations

27.771 Pilot compartment.
27.773 Pilot compartment view.
27.775 Windshields and windows.
27.777 Cockpit controls.
27.779 Motion and effect of cockpit controls.
27.783 Doors.
27.785 Seats, berths, litters, safety belts, and harnesses.
27.787 Cargo and baggage compartments.
27.801 Ditching.
27.805 Flight crew emergency exits.
27.807 Emergency exits.
27.831 Ventilation.
27.833 Heaters.

                             Fire Protection

27.853 Compartment interiors.
27.855 Cargo and baggage compartments.
27.859 Heating systems.
27.861 Fire protection of structure, controls, and other parts.

[[Page 659]]

27.863 Flammable fluid fire protection.

                             External Loads

27.865 External loads.

                              Miscellaneous

27.871 Leveling marks.
27.873 Ballast provisions.

                          Subpart E_Powerplant

                                 General

27.901 Installation.
27.903 Engines.
27.907 Engine vibration.

                           Rotor Drive System

27.917 Design.
27.921 Rotor brake.
27.923 Rotor drive system and control mechanism tests.
27.927 Additional tests.
27.931 Shafting critical speed.
27.935 Shafting joints.
27.939 Turbine engine operating characteristics.

                               Fuel System

27.951 General.
27.952 Fuel system crash resistance.
27.953 Fuel system independence.
27.954 Fuel system lightning protection.
27.955 Fuel flow.
27.959 Unusable fuel supply.
27.961 Fuel system hot weather operation.
27.963 Fuel tanks: general.
27.965 Fuel tank tests.
27.967 Fuel tank installation.
27.969 Fuel tank expansion space.
27.971 Fuel tank sump.
27.973 Fuel tank filler connection.
27.975 Fuel tank vents.
27.977 Fuel tank outlet.

                         Fuel System Components

27.991 Fuel pumps.
27.993 Fuel system lines and fittings.
27.995 Fuel valves.
27.997 Fuel strainer or filter.
27.999 Fuel system drains.

                               Oil System

27.1011 Engines: General.
27.1013 Oil tanks.
27.1015 Oil tank tests.
27.1017 Oil lines and fittings.
27.1019 Oil strainer or filter.
27.1021 Oil system drains.
27.1027 Transmissions and gearboxes: General.

                                 Cooling

27.1041 General.
27.1043 Cooling tests.
27.1045 Cooling test procedures.

                            Induction System

27.1091 Air induction.
27.1093 Induction system icing protection.

                             Exhaust System

27.1121 General.
27.1123 Exhaust piping.

                   Powerplant Controls and Accessories

27.1141 Powerplant controls: general.
27.1143 Engine controls.
27.1145 Ignition switches.
27.1147 Mixture controls.
27.1151 Rotor brake controls.
27.1163 Powerplant accessories.

                       Powerplant Fire Protection

27.1183 Lines, fittings, and components.
27.1185 Flammable fluids.
27.1187 Ventilation and drainage.
27.1189 Shutoff means.
27.1191 Firewalls.
27.1193 Cowling and engine compartment covering.
27.1194 Other surfaces.
27.1195 Fire detector systems.

                           Subpart F_Equipment

                                 General

27.1301 Function and installation.
27.1303 Flight and navigation instruments.
27.1305 Powerplant instruments.
27.1307 Miscellaneous equipment.
27.1309 Equipment, systems, and installations.
27.1316 Electrical and electronic system lightning protection.
27.1317 High-intensity Radiated Fields (HIRF) Protection.

                        Instruments: Installation

27.1321 Arrangement and visibility.
27.1322 Warning, caution, and advisory lights.
27.1323 Airspeed indicating system.
27.1325 Static pressure systems.
27.1327 Magnetic direction indicator.
27.1329 Automatic pilot system.
27.1335 Flight director systems.
27.1337 Powerplant instruments.

                    Electrical Systems and Equipment

27.1351 General.
27.1353 Storage battery design and installation.
27.1357 Circuit protective devices.
27.1361 Master switch.
27.1365 Electric cables.
27.1367 Switches.

                                 Lights

27.1381 Instrument lights.
27.1383 Landing lights.
27.1385 Position light system installation.

[[Page 660]]

27.1387 Position light system dihedral angles.
27.1389 Position light distribution and intensities.
27.1391 Minimum intensities in the horizontal plane of forward and rear 
          position lights.
27.1393 Minimum intensities in any vertical plane of forward and rear 
          position lights.
27.1395 Maximum intensities in overlapping beams of forward and rear 
          position lights.
27.1397 Color specifications.
27.1399 Riding light.
27.1401 Anticollision light system.

                            Safety Equipment

27.1411 General.
27.1413 Safety belts.
27.1415 Ditching equipment.
27.1419 Ice protection.
27.1435 Hydraulic systems.
27.1457 Cockpit voice recorders.
27.1459 Flight data recorders.
27.1461 Equipment containing high energy rotors.

             Subpart G_Operating Limitations and Information

27.1501 General.

                          Operating Limitations

27.1503 Airspeed limitations: general.
27.1505 Never-exceed speed.
27.1509 Rotor speed.
27.1519 Weight and center of gravity.
27.1521 Powerplant limitations.
27.1523 Minimum flight crew.
27.1525 Kinds of operations.
27.1527 Maximum operating altitude.
27.1529 Instructions for Continued Airworthiness.

                          Markings and Placards

27.1541 General.
27.1543 Instrument markings: general.
27.1545 Airspeed indicator.
27.1547 Magnetic direction indicator.
27.1549 Powerplant instruments.
27.1551 Oil quantity indicator.
27.1553 Fuel quantity indicator.
27.1555 Control markings.
27.1557 Miscellaneous markings and placards.
27.1559 Limitations placard.
27.1561 Safety equipment.
27.1565 Tail rotor.

          Rotorcraft Flight Manual and Approved Manual Material

27.1581 General.
27.1583 Operating limitations.
27.1585 Operating procedures.
27.1587 Performance information.
27.1589 Loading information.

Appendix A to Part 27--Instructions for Continued Airworthiness
Appendix B to Part 27--Airworthiness Criteria for Helicopter Instrument 
          Flight
Appendix C to Part 27--Criteria for Category A
Appendix D to Part 27--HIRF Environments and Equipment HIRF Test Levels

    Authority: 49 U.S.C. 106(g), 40113, 44701-44702, 44704.

    Source: Docket No. 5074, 29 FR 15695, Nov. 24, 1964, unless 
otherwise noted.



                            Subpart A_General



Sec. 27.1  Applicability.

    (a) This part prescribes airworthiness standards for the issue of 
type certificates, and changes to those certificates, for normal 
category rotorcraft with maximum weights of 7,000 pounds or less and 
nine or less passenger seats.
    (b) Each person who applies under Part 21 for such a certificate or 
change must show compliance with the applicable requirements of this 
part.
    (c) Multiengine rotorcraft may be type certified as Category A 
provided the requirements referenced in appendix C of this part are met.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-33, 
61 FR 21906, May 10, 1996; Amdt. 27-37, 64 FR 45094, Aug. 18, 1999]



Sec. 27.2  Special retroactive requirements.

    (a) For each rotorcraft manufactured after September 16, 1992, each 
applicant must show that each occupant's seat is equipped with a safety 
belt and shoulder harness that meets the requirements of paragraphs (a), 
(b), and (c) of this section.
    (1) Each occupant's seat must have a combined safety belt and 
shoulder harness with a single-point release. Each pilot's combined 
safety belt and shoulder harness must allow each pilot, when seated with 
safety belt and shoulder harness fastened, to perform all functions 
necessary for flight operations. There must be a means to secure belts 
and harnesses, when not in use, to prevent interference with the 
operation of the rotorcraft and with rapid egress in an emergency.

[[Page 661]]

    (2) Each occupant must be protected from serious head injury by a 
safety belt plus a shoulder harness that will prevent the head from 
contacting any injurious object.
    (3) The safety belt and shoulder harness must meet the static and 
dynamic strength requirements, if applicable, specified by the 
rotorcraft type certification basis.
    (4) For purposes of this section, the date of manufacture is 
either--
    (i) The date the inspection acceptance records, or equivalent, 
reflect that the rotorcraft is complete and meets the FAA-Approved Type 
Design Data; or
    (ii) The date the foreign civil airworthiness authority certifies 
that the rotorcraft is complete and issues an original standard 
airworthiness certificate, or equivalent, in that country.
    (b) For rotorcraft with a certification basis established prior to 
October 18, 1999--
    (1) The maximum passenger seat capacity may be increased to eight or 
nine provided the applicant shows compliance with all the airworthiness 
requirements of this part in effect on October 18, 1999.
    (2) The maximum weight may be increased to greater than 6,000 pounds 
provided--
    (i) The number of passenger seats is not increased above the maximum 
number certificated on October 18, 1999, or
    (ii) The applicant shows compliance with all of the airworthiness 
requirements of this part in effect on October 18, 1999.

[Doc. No. 26078, 56 FR 41051, Aug. 16, 1991, as amended by Amdt. 27-37, 
64 FR 45094, Aug. 18, 1999]



                            Subpart B_Flight

                                 General



Sec. 27.21  Proof of compliance.

    Each requirement of this subpart must be met at each appropriate 
combination of weight and center of gravity within the range of loading 
conditions for which certification is requested. This must be shown--
    (a) By tests upon a rotorcraft of the type for which certification 
is requested, or by calculations based on, and equal in accuracy to, the 
results of testing; and
    (b) By systematic investigation of each required combination of 
weight and center of gravity if compliance cannot be reasonably inferred 
from combinations investigated.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 
49 FR 44432, Nov. 6, 1984]



Sec. 27.25  Weight limits.

    (a) Maximum weight. The maximum weight (the highest weight at which 
compliance with each applicable requirement of this part is shown) must 
be established so that it is--
    (1) Not more than--
    (i) The highest weight selected by the applicant;
    (ii) The design maximum weight (the highest weight at which 
compliance with each applicable structural loading condition of this 
part is shown);
    (iii) The highest weight at which compliance with each applicable 
flight requirement of this part is shown; or
    (iv) The highest weight in which the provisions of Sec. Sec. 27.87 
or 27.143(c)(1), or combinations thereof, are demonstrated if the 
weights and operating conditions (altitude and temperature) prescribed 
by those requirements cannot be met; and
    (2) Not less than the sum of--
    (i) The empty weight determined under Sec. 27.29; and
    (ii) The weight of usable fuel appropriate to the intended operation 
with full payload;
    (iii) The weight of full oil capacity; and
    (iv) For each seat, an occupant weight of 170 pounds or any lower 
weight for which certification is requested.
    (b) Minimum weight. The minimum weight (the lowest weight at which 
compliance with each applicable requirement of this part is shown) must 
be established so that it is--
    (1) Not more than the sum of--
    (i) The empty weight determined under Sec. 27.29; and
    (ii) The weight of the minimum crew necessary to operate the 
rotorcraft, assuming for each crewmember a weight no more than 170 
pounds, or any lower

[[Page 662]]

weight selected by the applicant or included in the loading 
instructions; and
    (2) Not less than--
    (i) The lowest weight selected by the applicant;
    (ii) The design minimum weight (the lowest weight at which 
compliance with each applicable structural loading condition of this 
part is shown); or
    (iii) The lowest weight at which compliance with each applicable 
flight requirement of this part is shown.
    (c) Total weight with jettisonable external load. A total weight for 
the rotorcraft with a jettisonable external load attached that is 
greater than the maximum weight established under paragraph (a) of this 
section may be established for any rotorcraft-load combination if--
    (1) The rotorcraft-load combination does not include human external 
cargo,
    (2) Structural component approval for external load operations under 
either Sec. 27.865 or under equivalent operational standards is 
obtained,
    (3) The portion of the total weight that is greater than the maximum 
weight established under paragraph (a) of this section is made up only 
of the weight of all or part of the jettisonable external load,
    (4) Structural components of the rotorcraft are shown to comply with 
the applicable structural requirements of this part under the increased 
loads and stresses caused by the weight increase over that established 
under paragraph (a) of this section, and
    (5) Operation of the rotorcraft at a total weight greater than the 
maximum certificated weight established under paragraph (a) of this 
section is limited by appropriate operating limitations under Sec. 
27.865(a) and (d) of this part.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 29, 1964, as amended by Amdt. 27-11, 
41 FR 55468, Dec. 20, 1976; Amdt. 25-42, 43 FR 2324, Jan. 16, 1978; 
Amdt. 27-36, 64 FR 43019, Aug. 6, 1999; Amdt. 27-44, 73 FR 10998, Feb. 
29, 2008; 73 FR 33876, June 16, 2008]



Sec. 27.27  Center of gravity limits.

    The extreme forward and aft centers of gravity and, where critical, 
the extreme lateral centers of gravity must be established for each 
weight established under Sec. 27.25. Such an extreme may not lie 
beyond--
    (a) The extremes selected by the applicant;
    (b) The extremes within which the structure is proven; or
    (c) The extremes within which compliance with the applicable flight 
requirements is shown.

[Amdt. 27-2, 33 FR 962, Jan. 26, 1968]



Sec. 27.29  Empty weight and corresponding center of gravity.

    (a) The empty weight and corresponding center of gravity must be 
determined by weighing the rotorcraft without the crew and payload, but 
with--
    (1) Fixed ballast;
    (2) Unusable fuel; and
    (3) Full operating fluids, including--
    (i) Oil;
    (ii) Hydraulic fluid; and
    (iii) Other fluids required for normal operation of roto-craft 
systems, except water intended for injection in the engines.
    (b) The condition of the rotorcraft at the time of determining empty 
weight must be one that is well defined and can be easily repeated, 
particularly with respect to the weights of fuel, oil, coolant, and 
installed equipment.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 
43 FR 2324, Jan. 16, 1978]



Sec. 27.31  Removable ballast.

    Removable ballast may be used in showing compliance with the flight 
requirements of this subpart.



Sec. 27.33  Main rotor speed and pitch limits.

    (a) Main rotor speed limits. A range of main rotor speeds must be 
established that--

[[Page 663]]

    (1) With power on, provides adequate margin to accommodate the 
variations in rotor speed occurring in any appropriate maneuver, and is 
consistent with the kind of governor or synchronizer used; and
    (2) With power off, allows each appropriate autorotative maneuver to 
be performed throughout the ranges of airspeed and weight for which 
certification is requested.
    (b) Normal main rotor high pitch limits (power on). For rotocraft, 
except helicopters required to have a main rotor low speed warning under 
paragraph (e) of this section. It must be shown, with power on and 
without exceeding approved engine maximum limitations, that main rotor 
speeds substantially less than the minimum approved main rotor speed 
will not occur under any sustained flight condition. This must be met 
by--
    (1) Appropriate setting of the main rotor high pitch stop;
    (2) Inherent rotorcraft characteristics that make unsafe low main 
rotor speeds unlikely; or
    (3) Adequate means to warn the pilot of unsafe main rotor speeds.
    (c) Normal main rotor low pitch limits (power off). It must be 
shown, with power off, that--
    (1) The normal main rotor low pitch limit provides sufficient rotor 
speed, in any autorotative condition, under the most critical 
combinations of weight and airspeed; and
    (2) It is possible to prevent overspeeding of the rotor without 
exceptional piloting skill.
    (d) Emergency high pitch. If the main rotor high pitch stop is set 
to meet paragraph (b)(1) of this section, and if that stop cannot be 
exceeded inadvertently, additional pitch may be made available for 
emergency use.
    (e) Main rotor low speed warning for helicopters. For each single 
engine helicopter, and each multiengine helicopter that does not have an 
approved device that automatically increases power on the operating 
engines when one engine fails, there must be a main rotor low speed 
warning which meets the following requirements:
    (1) The warning must be furnished to the pilot in all flight 
conditions, including power-on and power-off flight, when the speed of a 
main rotor approaches a value that can jeopardize safe flight.
    (2) The warning may be furnished either through the inherent 
aerodynamic qualities of the helicopter or by a device.
    (3) The warning must be clear and distinct under all conditons, and 
must be clearly distinguishable from all other warnings. A visual device 
that requires the attention of the crew within the cockpit is not 
acceptable by itself.
    (4) If a warning device is used, the device must automatically 
deactivate and reset when the low-speed condition is corrected. If the 
device has an audible warning, it must also be equipped with a means for 
the pilot to manually silence the audible warning before the low-speed 
condition is corrected.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 962, Jan. 26, 1968; Amdt. 27-14, 43 FR 2324, Jan. 16, 1978]

                               Performance



Sec. 27.45  General.

    (a) Unless otherwise prescribed, the performance requirements of 
this subpart must be met for still air and a standard atmosphere.
    (b) The performance must correspond to the engine power available 
under the particular ambient atmospheric conditions, the particular 
flight condition, and the relative humidity specified in paragraphs (d) 
or (e) of this section, as appropriate.
    (c) The available power must correspond to engine power, not 
exceeding the approved power, less--
    (1) Installation losses; and
    (2) The power absorbed by the accessories and services appropriate 
to the particular ambient atmopheric conditions and the particular 
flight condition.
    (d) For reciprocating engine-powered rotorcraft, the performance, as 
affected by engine power, must be based on a relative humidity of 80 
percent in a standard atmosphere.

[[Page 664]]

    (e) For turbine engine-powered rotorcraft, the performance, as 
affected by engine power, must be based on a relative humidity of--
    (1) 80 percent, at and below standard temperature; and
    (2) 34 percent, at and above standard temperature plus 50 degrees F. 
Between these two temperatures, the relative humidity must vary 
linearly.
    (f) For turbine-engine-powered rotorcraft, a means must be provided 
to permit the pilot to determine prior to takeoff that each engine is 
capable of developing the power necessary to achieve the applicable 
rotorcraft performance prescribed in this subpart.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 27-14, 43 FR 2324, Jan. 16, 1978, as amended by Amdt. 27-21, 49 
FR 44432, Nov. 6, 1984]



Sec. 27.49  Performance at minimum operating speed.

    (a) For helicopters--
    (1) The hovering ceiling must be determined over the ranges of 
weight, altitude, and temperature for which certification is requested, 
with--
    (i) Takeoff power;
    (ii) The landing gear extended; and
    (iii) The helicopter in-ground effect at a height consistent with 
normal takeoff procedures; and
    (2) The hovering ceiling determined under paragraph (a)(1) of this 
section must be at least--
    (i) For reciprocating engine powered helicopters, 4,000 feet at 
maximum weight with a standard atmosphere;
    (ii) For turbine engine powered helicopters, 2,500 feet pressure 
altitude at maximum weight at a temperature of standard plus 22 [deg]C 
(standard plus 40 [deg]F).
    (3) The out-of-ground effect hovering performance must be determined 
over the ranges of weight, altitude, and temperature for which 
certification is requested, using takeoff power.
    (b) For rotorcraft other than helicopters, the steady rate of climb 
at the minimum operating speed must be determined over the ranges of 
weight, altitude, and temperature for which certification is requested, 
with--
    (1) Takeoff power; and
    (2) The landing gear extended.

[Amdt. 27-44, 73 FR 10998, Feb. 29, 2008]



Sec. 27.51  Takeoff.

    The takeoff, with takeoff power and r.p.m. at the most critical 
center of gravity, and with weight from the maximum weight at sea level 
to the weight for which takeoff certification is requested for each 
altitude covered by this section--
    (a) May not require exceptional piloting skill or exceptionally 
favorable conditions throughout the ranges of altitude from standard sea 
level conditions to the maximum altitude for which takeoff and landing 
certification is requested, and
    (b) Must be made in such a manner that a landing can be made safely 
at any point along the flight path if an engine fails. This must be 
demonstrated up to the maximum altitude for which takeoff and landing 
certification is requested or 7,000 feet density altitude, whichever is 
less.

[Amdt. 27-44, 73 FR 10999, Feb. 29, 2008]



Sec. 27.65  Climb: all engines operating.

    (a) For rotorcraft other than helicopters--
    (1) The steady rate of climb, at VY, must be determined--
    (i) With maximum continuous power on each engine;
    (ii) With the landing gear retracted; and
    (iii) For the weights, altitudes, and temperatures for which 
certification is requested; and
    (2) The climb gradient, at the rate of climb determined in 
accordance with paragraph (a)(1) of this section, must be either--
    (i) At least 1:10 if the horizontal distance required to take off 
and climb over a 50-foot obstacle is determined for each weight, 
altitude, and temperature within the range for which certification is 
requested; or
    (ii) At least 1:6 under standard sea level conditions.
    (b) Each helicopter must meet the following requirements:
    (1) VY must be determined--

[[Page 665]]

    (i) For standard sea level conditions;
    (ii) At maximum weight; and
    (iii) With maximum continuous power on each engine.
    (2) The steady rate of climb must be determined--
    (i) At the climb speed selected by the applicant at or below VNE;
    (ii) Within the range from sea level up to the maximum altitude for 
which certification is requested;
    (iii) For the weights and temperatures that correspond to the 
altitude range set forth in paragraph (b)(2)(ii) of this section and for 
which certification is requested; and
    (iv) With maximum continuous power on each engine.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 
43 FR 2324, Jan. 16, 1978; Amdt. 27-33, 61 FR 21907, May 10, 1996]



Sec. 27.67  Climb: one engine inoperative.

    For multiengine helicopters, the steady rate of climb (or descent), 
at Vy (or at the speed for minimum rate of descent), must be determined 
with--
    (a) Maximum weight;
    (b) The critical engine inoperative and the remaining engines at 
either--
    (1) Maximum continuous power and, for helicopters for which 
certification for the use of 30-minute OEI power is requested, at 30-
minute OEI power; or
    (2) Continuous OEI power for helicopters for which certification for 
the use of continuous OEI power is requested.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 53 
FR 34210, Sept. 2, 1988]



Sec. 27.71  Autorotation performance.

    For single-engine helicopters and multiengine helicopters that do 
not meet the Category A engine isolation requirements of Part 29 of this 
chapter, the minimum rate of descent airspeed and the best angle-of-
glide airspeed must be determined in autorotation at--
    (a) Maximum weight; and
    (b) Rotor speed(s) selected by the applicant.

[Amdt. 27-21, 49 FR 44433, Nov. 6, 1984]



Sec. 27.75  Landing.

    (a) The rotorcraft must be able to be landed with no excessive 
vertical acceleration, no tendency to bounce, nose over, ground loop, 
porpoise, or water loop, and without exceptional piloting skill or 
exceptionally favorable conditions, with--
    (1) Approach or autorotation speeds appropriate to the type of 
rotorcraft and selected by the applicant;
    (2) The approach and landing made with--
    (i) Power off, for single engine rotorcraft and entered from steady 
state autorotation; or
    (ii) One-engine inoperative (OEI) for multiengine rotorcraft, with 
each operating engine within approved operating limitations, and entered 
from an established OEI approach.
    (b) Multiengine rotorcraft must be able to be landed safely after 
complete power failure under normal operating conditions.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 
43 FR 2324, Jan. 16, 1978; Amdt. 27-44, 73 FR 10999, Feb. 29, 2008]



Sec. 27.87  Height-speed envelope.

    (a) If there is any combination of height and forward speed 
(including hover) under which a safe landing cannot be made under the 
applicable power failure condition in paragraph (b) of this section, a 
limiting height-speed envelope must be established (including all 
pertinent information) for that condition, throughout the ranges of--
    (1) Altitude, from standard sea level conditions to the maximum 
altitude capability of the rotorcraft, or 7000 feet density altitude, 
whichever is less; and
    (2) Weight, from the maximum weight at sea level to the weight 
selected by the applicant for each altitude covered by paragraph (a)(1) 
of this section. For helicopters, the weight at altitudes above sea 
level may not be less than the maximum weight or the

[[Page 666]]

highest weight allowing hovering out-of-ground effect, whichever is 
lower.
    (b) The applicable power failure conditions are--
    (1) For single-engine helicopters, full autorotation;
    (2) For multiengine helicopters, OEI (where engine isolation 
features ensure continued operation of the remaining engines), and the 
remaining engine(s) within approved limits and at the minimum installed 
specification power available for the most critical combination of 
approved ambient temperature and pressure altitude resulting in 7000 
feet density altitude or the maximum altitude capability of the 
helicopter, whichever is less, and
    (3) For other rotorcraft, conditions appropriate to the type.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 
43 FR 2324, Jan. 16, 1978; Amdt. 27-21, 49 FR 44433, Nov. 6, 1984; Amdt. 
27-44, 73 FR 10999, Feb. 29, 2008]

                         Flight Characteristics



Sec. 27.141  General.

    The rotorcraft must--
    (a) Except as specifically required in the applicable section, meet 
the flight characteristics requirements of this subpart--
    (1) At the altitudes and temperatures expected in operation;
    (2) Under any critical loading condition within the range of weights 
and centers of gravity for which certification is requested;
    (3) For power-on operations, under any condition of speed, power, 
and rotor r.p.m. for which certification is requested; and
    (4) For power-off operations, under any condition of speed and rotor 
r.p.m. for which certification is requested that is attainable with the 
controls rigged in accordance with the approved rigging instructions and 
tolerances;
    (b) Be able to maintain any required flight condition and make a 
smooth transition from any flight condition to any other flight 
condition without exceptional piloting skill, alertness, or strength, 
and without danger of exceeding the limit load factor under any 
operating condition probable for the type, including--
    (1) Sudden failure of one engine, for multiengine rotorcraft meeting 
Transport Category A engine isolation requirements of Part 29 of this 
chapter;
    (2) Sudden, complete power failure for other rotorcraft; and
    (3) Sudden, complete control system failures specified in Sec. 
27.695 of this part; and
    (c) Have any additional characteristic required for night or 
instrument operation, if certification for those kinds of operation is 
requested. Requirements for helicopter instrument flight are contained 
in appendix B of this part.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 962, Jan. 26, 1968; Amdt. 27-11, 41 FR 55468, Dec. 20, 1976; Amdt. 
27-19, 48 FR 4389, Jan. 31, 1983; Amdt. 27-21, 49 FR 44433, Nov. 6, 
1984]



Sec. 27.143  Controllability and maneuverability.

    (a) The rotorcraft must be safely controllable and maneuverable--
    (1) During steady flight; and
    (2) During any maneuver appropriate to the type, including--
    (i) Takeoff;
    (ii) Climb;
    (iii) Level flight;
    (iv) Turning flight;
    (v) Autorotation;
    (vi) Landing (power on and power off); and
    (vii) Recovery to power-on flight from a balked autorotative 
approach.
    (b) The margin of cyclic control must allow satisfactory roll and 
pitch control at VNE with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Critical rotor r.p.m.; and
    (4) Power off (except for helicopters demonstrating compliance with 
paragraph (f) of this section) and power on.
    (c) Wind velocities from zero to at least 17 knots, from all 
azimuths, must be established in which the rotorcraft can be operated 
without loss of control on or near the ground in any maneuver 
appropriate to the type (such as crosswind takeoffs, sideward flight, 
and rearward flight)--

[[Page 667]]

    (1) With altitude, from standard sea level conditions to the maximum 
takeoff and landing altitude capability of the rotorcraft or 7000 feet 
density altitude, whichever is less; with--
    (i) Critical Weight;
    (ii) Critical center of gravity;
    (iii) Critical rotor r.p.m.;
    (2) For takeoff and landing altitudes above 7000 feet density 
altitude with--
    (i) Weight selected by the applicant;
    (ii) Critical center of gravity; and
    (iii) Critical rotor r.p.m.
    (d) Wind velocities from zero to at least 17 knots, from all 
azimuths, must be established in which the rotorcraft can be operated 
without loss of control out-of-ground-effect, with--
    (1) Weight selected by the applicant;
    (2) Critical center of gravity;
    (3) Rotor r.p.m. selected by the applicant; and
    (4) Altitude, from standard sea level conditions to the maximum 
takeoff and landing altitude capability of the rotorcraft.
    (e) The rotorcraft, after (1) failure of one engine in the case of 
multiengine rotorcraft that meet Transport Category A engine isolation 
requirements, or (2) complete engine failure in the case of other 
rotorcraft, must be controllable over the range of speeds and altitudes 
for which certification is requested when such power failure occurs with 
maximum continuous power and critical weight. No corrective action time 
delay for any condition following power failure may be less than--
    (i) For the cruise condition, one second, or normal pilot reaction 
time (whichever is greater); and
    (ii) For any other condition, normal pilot reaction time.
    (f) For helicopters for which a VNE (power-off) is established under 
Sec. 27.1505(c), compliance must be demonstrated with the following 
requirements with critical weight, critical center of gravity, and 
critical rotor r.p.m.:
    (1) The helicopter must be safely slowed to VNE (power-off), without 
exceptional pilot skill, after the last operating engine is made 
inoperative at power-on VNE.
    (2) At a speed of 1.1 VNE (power-off), the margin of cyclic control 
must allow satisfactory roll and pitch control with power off.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 963, Jan. 26, 1968; Amdt. 27-14, 43 FR 2325, Jan. 16, 1978; Amdt. 27-
21, 49 FR 44433, Nov. 6, 1984; Amdt. 27-44, 73 FR 10999, Feb. 29, 2008]



Sec. 27.151  Flight controls.

    (a) Longitudinal, lateral, directional, and collective controls may 
not exhibit excessive breakout force, friction, or preload.
    (b) Control system forces and free play may not inhibit a smooth, 
direct rotorcraft response to control system input.

[Amdt. 27-21, 49 FR 44433, Nov. 6, 1984]



Sec. 27.161  Trim control.

    The trim control--
    (a) Must trim any steady longitudinal, lateral, and collective 
control forces to zero in level flight at any appropriate speed; and
    (b) May not introduce any undesirable discontinuities in control 
force gradients.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 
49 FR 44433, Nov. 6, 1984]



Sec. 27.171  Stability: general.

    The rotorcraft must be able to be flown, without undue pilot fatigue 
or strain, in any normal maneuver for a period of time as long as that 
expected in normal operation. At least three landings and takeoffs must 
be made during this demonstration.



Sec. 27.173  Static longitudinal stability.

    (a) The longitudinal control must be designed so that a rearward 
movement of the control is necessary to obtain an airspeed less than the 
trim speed, and a forward movement of the control is necessary to obtain 
an airspeed more than the trim speed.

[[Page 668]]

    (b) Throughout the full range of altitude for which certification is 
requested, with the throttle and collective pitch held constant during 
the maneuvers specified in Sec. 27.175(a) through (d), the slope of the 
control position versus airspeed curve must be positive. However, in 
limited flight conditions or modes of operation determined by the 
Administrator to be acceptable, the slope of the control position versus 
airspeed curve may be neutral or negative if the rotorcraft possesses 
flight characteristics that allow the pilot to maintain airspeed within 
5 knots of the desired trim airspeed without 
exceptional piloting skill or alertness.

[Amdt. 27-21, 49 FR 44433, Nov. 6, 1984, as amended by Amdt. 27-44, 73 
FR 10999, Feb. 29, 2008]



Sec. 27.175  Demonstration of static longitudinal stability.

    (a) Climb. Static longitudinal stability must be shown in the climb 
condition at speeds from Vy - 10 kt to Vy + 10 kt with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Maximum continuous power;
    (4) The landing gear retracted; and
    (5) The rotorcraft trimmed at VY.
    (b) Cruise. Static longitudinal stability must be shown in the 
cruise condition at speeds from 0.8 VNE - 10 kt to 0.8 
VNE + 10 kt or, if VH is less than 0.8 
VNE, from VH -10 kt to VH + 10 kt, 
with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Power for level flight at 0.8 VNE or VH, 
whichever is less;
    (4) The landing gear retracted; and
    (5) The rotorcraft trimmed at 0.8 VNE or VH, 
whichever is less.
    (c) VNE. Static longitudinal stability must be shown at 
speeds from VNE - 20 kt to VNE with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Power required for level flight at VNE -10 kt or 
maximum continuous power, whichever is less;
    (4) The landing gear retracted; and
    (5) The rotorcraft trimmed at VNE - 10 kt.
    (d) Autorotation. Static longitudinal stability must be shown in 
autorotation at--
    (1) Airspeeds from the minimum rate of descent airspeed-10 kt to the 
minimum rate of descent airspeed + 10 kt, with--
    (i) Critical weight;
    (ii) Critical center of gravity;
    (iii) The landing gear extended; and
    (iv) The rotorcraft trimmed at the minimum rate of descent airspeed.
    (2) Airspeeds from best angle-of-glide airspeed-10 kt to the best 
angle-of-glide airspeed + 10 kt, with--
    (i) Critical weight;
    (ii) Critical center of gravity;
    (iii) The landing gear retracted; and
    (iv) The rotorcraft trimmed at the best angle-of-glide airspeed.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 963, Jan. 26, 1968; Amdt. 27-11, 41 FR 55468, Dec. 20, 1976; Amdt. 
27-14, 43 FR 2325, Jan. 16, 1978; Amdt. 27-21, 49 FR 44433, Nov. 6, 
1984; Amdt. 27-34, 62 FR 46173, Aug. 29, 1997; Amdt. 27-44, 73 FR 10999, 
Feb. 29, 2008]



Sec. 27.177  Static directional stability.

    (a) The directional controls must operate in such a manner that the 
sense and direction of motion of the rotorcraft following control 
displacement are in the direction of the pedal motion with the throttle 
and collective controls held constant at the trim conditions specified 
in Sec. 27.175(a), (b), and (c). Sideslip angles must increase with 
steadily increasing directional control deflection for sideslip angles 
up to the lesser of--
    (1) 25 degrees from trim at a speed of 15 
knots less than the speed for minimum rate of descent varying linearly 
to 10 degrees from trim at VNE;
    (2) The steady state sideslip angles established by Sec. 27.351;
    (3) A sideslip angle selected by the applicant, which corresponds to 
a sideforce of at least 0.1g; or
    (4) The sideslip angle attained by maximum directional control 
input.
    (b) Sufficient cues must accompany the sideslip to alert the pilot 
when the aircraft is approaching the sideslip limits.

[[Page 669]]

    (c) During the maneuver specified in paragraph (a) of this section, 
the sideslip angle versus directional control position curve may have a 
negative slope within a small range of angles around trim, provided the 
desired heading can be maintained without exceptional piloting skill or 
alertness.

[Amdt. 27-44, 73 FR 11000, Feb. 29, 2008]

                Ground and Water Handling Characteristics



Sec. 27.231  General.

    The rotorcraft must have satisfactory ground and water handling 
characteristics, including freedom from uncontrollable tendencies in any 
condition expected in operation.



Sec. 27.235  Taxiing condition.

    The rotorcraft must be designed to withstand the loads that would 
occur when the rotorcraft is taxied over the roughest ground that may 
reasonably be expected in normal operation.



Sec. 27.239  Spray characteristics.

    If certification for water operation is requested, no spray 
characteristics during taxiing, takeoff, or landing may obscure the 
vision of the pilot or damage the rotors, propellers, or other parts of 
the rotorcraft.



Sec. 27.241  Ground resonance.

    The rotorcraft may have no dangerous tendency to oscillate on the 
ground with the rotor turning.

                    Miscellaneous Flight Requirements



Sec. 27.251  Vibration.

    Each part of the rotorcraft must be free from excessive vibration 
under each appropriate speed and power condition.



                     Subpart C_Strength Requirements

                                 General



Sec. 27.301  Loads.

    (a) Strength requirements are specified in terms of limit loads (the 
maximum loads to be expected in service) and ultimate loads (limit loads 
multiplied by prescribed factors of safety). Unless otherwise provided, 
prescribed loads are limit loads.
    (b) Unless otherwise provided, the specified air, ground, and water 
loads must be placed in equilibrium with inertia forces, considering 
each item of mass in the rotorcraft. These loads must be distributed to 
closely approximate or conservatively represent actual conditions.
    (c) If deflections under load would significantly change the 
distribution of external or internal loads, this redistribution must be 
taken into account.



Sec. 27.303  Factor of safety.

    Unless otherwise provided, a factor of safety of 1.5 must be used. 
This factor applies to external and inertia loads unless its application 
to the resulting internal stresses is more conservative.



Sec. 27.305  Strength and deformation.

    (a) The structure must be able to support limit loads without 
detrimental or permanent deformation. At any load up to limit loads, the 
deformation may not interfere with safe operation.
    (b) The structure must be able to support ultimate loads without 
failure. This must be shown by--
    (1) Applying ultimate loads to the structure in a static test for at 
least three seconds; or
    (2) Dynamic tests simulating actual load application.



Sec. 27.307  Proof of structure.

    (a) Compliance with the strength and deformation requirements of 
this subpart must be shown for each critical loading condition 
accounting for the environment to which the structure will be exposed in 
operation. Structural analysis (static or fatigue) may be used only if 
the structure conforms to those structures for which experience has 
shown this method to be reliable. In other cases, substantiating load 
tests must be made.
    (b) Proof of compliance with the strength requirements of this 
subpart must include--
    (1) Dynamic and endurance tests of rotors, rotor drives, and rotor 
controls;
    (2) Limit load tests of the control system, including control 
surfaces;

[[Page 670]]

    (3) Operation tests of the control system;
    (4) Flight stress measurement tests;
    (5) Landing gear drop tests; and
    (6) Any additional test required for new or unusual design features.

(Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, 1425)

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-3, 33 
FR 14105, Sept. 18, 1968; Amdt. 27-26, 55 FR 7999, Mar. 6, 1990]



Sec. 27.309  Design limitations.

    The following values and limitations must be established to show 
compliance with the structural requirements of this subpart:
    (a) The design maximum weight.
    (b) The main rotor r.p.m. ranges power on and power off.
    (c) The maximum forward speeds for each main rotor r.p.m. within the 
ranges determined under paragraph (b) of this section.
    (d) The maximum rearward and sideward flight speeds.
    (e) The center of gravity limits corresponding to the limitations 
determined under paragraphs (b), (c), and (d) of this section.
    (f) The rotational speed ratios between each powerplant and each 
connected rotating component.
    (g) The positive and negative limit maneuvering load factors.

                              Flight Loads



Sec. 27.321  General.

    (a) The flight load factor must be assumed to act normal to the 
longitudinal axis of the rotorcraft, and to be equal in magnitude and 
opposite in direction to the rotorcraft inertia load factor at the 
center of gravity.
    (b) Compliance with the flight load requirements of this subpart 
must be shown--
    (1) At each weight from the design minimum weight to the design 
maximum weight; and
    (2) With any practical distribution of disposable load within the 
operating limitations in the Rotorcraft Flight Manual.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55468, Dec. 20, 1976]



Sec. 27.337  Limit maneuvering load factor.

    The rotorcraft must be designed for--
    (a) A limit maneuvering load factor ranging from a positive limit of 
3.5 to a negative limit of -1.0; or
    (b) Any positive limit maneuvering load factor not less than 2.0 and 
any negative limit maneuvering load factor of not less than -0.5 for 
which--
    (1) The probability of being exceeded is shown by analysis and 
flight tests to be extremely remote; and
    (2) The selected values are appropriate to each weight condition 
between the design maximum and design minimum weights.

[Amdt. 27-26, 55 FR 7999, Mar. 6, 1990]



Sec. 27.339  Resultant limit maneuvering loads.

    The loads resulting from the application of limit maneuvering load 
factors are assumed to act at the center of each rotor hub and at each 
auxiliary lifting surface, and to act in directions, and with 
distributions of load among the rotors and auxiliary lifting surfaces, 
so as to represent each critical maneuvering condition, including power-
on and power-off flight with the maximum design rotor tip speed ratio. 
The rotor tip speed ratio is the ratio of the rotorcraft flight velocity 
component in the plane of the rotor disc to the rotational tip speed of 
the rotor blades, and is expressed as follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.083

where--

V= The airspeed along flight path (f.p.s.);
a= The angle between the projection, in the plane of symmetry, of the 
          axis of no feathering and a line perpendicular to the flight 
          path (radians, positive when axis is pointing aft);
omega= The angular velocity of rotor (radians per second); and
R= The rotor radius (ft).

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55469, Dec. 20, 1976]



Sec. 27.341  Gust loads.

    The rotorcraft must be designed to withstand, at each critical 
airspeed including hovering, the loads resulting

[[Page 671]]

from a vertical gust of 30 feet per second.



Sec. 27.351  Yawing conditions.

    (a) Each rotorcraft must be designed for the loads resulting from 
the maneuvers specified in paragraphs (b) and (c) of this section with--
    (1) Unbalanced aerodynamic moments about the center of gravity which 
the aircraft reacts to in a rational or conservative manner considering 
the principal masses furnishing the reacting inertia forces; and
    (2) Maximum main rotor speed.
    (b) To produce the load required in paragraph (a) of this section, 
in unaccelerated flight with zero yaw, at forward speeds from zero up to 
0.6 VNE--
    (1) Displace the cockpit directional control suddenly to the maximum 
deflection limited by the control stops or by the maximum pilot force 
specified in Sec. 27.397(a);
    (2) Attain a resulting sideslip angle or 90[deg], whichever is less; 
and
    (3) Return the directional control suddenly to neutral.
    (c) To produce the load required in paragraph (a) of this section, 
in unaccelerated flight with zero yaw, at forward speeds from 0.6 
VNE up to VNE or VH, whichever is 
less--
    (1) Displace the cockpit directional control suddenly to the maximum 
deflection limited by the control stops or by the maximum pilot force 
specified in Sec. 27.397(a);
    (2) Attain a resulting sideslip angle or 15[deg], whichever is less, 
at the lesser speed of VNE or VH;
    (3) Vary the sideslip angles of paragraphs (b)(2) and (c)(2) of this 
section directly with speed; and
    (4) Return the directional control suddenly to neutral.

[Amdt. 27-26, 55 FR 7999, Mar. 6, 1990, as amended by Amdt. 27-34, 62 FR 
46173, Aug. 29, 1997]



Sec. 27.361  Engine torque.

    (a) For turbine engines, the limit torque may not be less than the 
highest of--
    (1) The mean torque for maximum continuous power multiplied by 1.25;
    (2) The torque required by Sec. 27.923;
    (3) The torque required by Sec. 27.927; or
    (4) The torque imposed by sudden engine stoppage due to malfunction 
or structural failure (such as compressor jamming).
    (b) For reciprocating engines, the limit torque may not be less than 
the mean torque for maximum continuous power multiplied by--
    (1) 1.33, for engines with five or more cylinders; and
    (2) Two, three, and four, for engines with four, three, and two 
cylinders, respectively.

[Amdt. 27-23, 53 FR 34210, Sept. 2, 1988]

                    Control Surface and System Loads



Sec. 27.391  General.

    Each auxiliary rotor, each fixed or movable stabilizing or control 
surface, and each system operating any flight control must meet the 
requirements of Sec. Sec. 27.395, 27.397, 27.399, 27.411, and 27.427.

[Amdt. 27-26, 55 FR 7999, Mar. 6, 1990, as amended by Amdt. 27-34, 62 FR 
46173, Aug. 29, 1997]



Sec. 27.395  Control system.

    (a) The part of each control system from the pilot's controls to the 
control stops must be designed to withstand pilot forces of not less 
than--
    (1) The forces specified in Sec. 27.397; or
    (2) If the system prevents the pilot from applying the limit pilot 
forces to the system, the maximum forces that the system allows the 
pilot to apply, but not less than 0.60 times the forces specified in 
Sec. 27.397.
    (b) Each primary control system, including its supporting structure, 
must be designed as follows:
    (1) The system must withstand loads resulting from the limit pilot 
forces prescribed in Sec. 27.397.
    (2) Notwithstanding paragraph (b)(3) of this section, when power-
operated actuator controls or power boost controls are used, the system 
must also withstand the loads resulting from the force output of each 
normally energized power device, including any single power boost or 
actuator system failure.
    (3) If the system design or the normal operating loads are such that 
a part of the system cannot react to the limit pilot forces prescribed 
in Sec. 27.397, that part of the system must be designed to

[[Page 672]]

withstand the maximum loads that can be obtained in normal operation. 
The minimum design loads must, in any case, provide a rugged system for 
service use, including consideration of fatigue, jamming, ground gusts, 
control inertia, and friction loads. In the absence of rational 
analysis, the design loads resulting from 0.60 of the specified limit 
pilot forces are acceptable minimum design loads.
    (4) If operational loads may be exceeded through jamming, ground 
gusts, control inertia, or friction, the system must withstand the limit 
pilot forces specified in Sec. 27.397, without yielding.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-26, 
55 FR 7999, Mar. 6, 1990]



Sec. 27.397  Limit pilot forces and torques.

    (a) Except as provided in paragraph (b) of this section, the limit 
pilot forces are as follows:
    (1) For foot controls, 130 pounds.
    (2) For stick controls, 100 pounds fore and aft, and 67 pounds 
laterally.
    (b) For flap, tab, stabilizer, rotor brake, and landing gear 
operating controls, the follows apply (R=radius in inches):
    (1) Crank, wheel, and lever controls, [1+R]/3 x 50 pounds, but not 
less than 50 pounds nor more than 100 pounds for hand operated controls 
or 130 pounds for foot operated controls, applied at any angle within 20 
degrees of the plane of motion of the control.
    (2) Twist controls, 80R inch-pounds.

[Amdt. 27-11, 41 FR 55469, Dec. 20, 1976, as amended by Amdt. 27-40, 66 
FR 23538, May 9, 2001]



Sec. 27.399  Dual control system.

    Each dual primary flight control system must be designed to 
withstand the loads that result when pilot forces of 0.75 times those 
obtained under Sec. 27.395 are applied--
    (a) In opposition; and
    (b) In the same direction.



Sec. 27.411  Ground clearance: tail rotor guard.

    (a) It must be impossible for the tail rotor to contact the landing 
surface during a normal landing.
    (b) If a tail rotor guard is required to show compliance with 
paragraph (a) of this section--
    (1) Suitable design loads must be established for the guard; and
    (2) The guard and its supporting structure must be designed to 
withstand those loads.



Sec. 27.427  Unsymmetrical loads.

    (a) Horizontal tail surfaces and their supporting structure must be 
designed for unsymmetrical loads arising from yawing and rotor wake 
effects in combination with the prescribed flight conditions.
    (b) To meet the design criteria of paragraph (a) of this section, in 
the absence of more rational data, both of the following must be met:
    (1) One hundred percent of the maximum loading from the symmetrical 
flight conditions acts on the surface on one side of the plane of 
symmetry, and no loading acts on the other side.
    (2) Fifty percent of the maximum loading from the symmetrical flight 
conditions acts on the surface on each side of the plane of symmetry but 
in opposite directions.
    (c) For empennage arrangements where the horizontal tail surfaces 
are supported by the vertical tail surfaces, the vertical tail surfaces 
and supporting structure must be designed for the combined vertical and 
horizontal surface loads resulting from each prescribed flight 
condition, considered separately. The flight conditions must be selected 
so the maximum design loads are obtained on each surface. In the absence 
of more rational data, the unsymmetrical horizontal tail surface loading 
distributions described in this section must be assumed.

[Admt. 27-26, 55 FR 7999, Mar. 6, 1990, as amended by Amdt. 27-27, 55 FR 
38966, Sept. 21, 1990]

                              Ground Loads



Sec. 27.471  General.

    (a) Loads and equilibrium. For limit ground loads--
    (1) The limit ground loads obtained in the landing conditions in 
this part must be considered to be external loads that would occur in 
the rotorcraft

[[Page 673]]

structure if it were acting as a rigid body; and
    (2) In each specified landing condition, the external loads must be 
placed in equilibrium with linear and angular inertia loads in a 
rational or conservative manner.
    (b) Critical centers of gravity. The critical centers of gravity 
within the range for which certification is requested must be selected 
so that the maximum design loads are obtained in each landing gear 
element.



Sec. 27.473  Ground loading conditions and assumptions.

    (a) For specified landing conditions, a design maximum weight must 
be used that is not less than the maximum weight. A rotor lift may be 
assumed to act through the center of gravity throughout the landing 
impact. This lift may not exceed two-thirds of the design maximum 
weight.
    (b) Unless otherwise prescribed, for each specified landing 
condition, the rotorcraft must be designed for a limit load factor of 
not less than the limit inertia load factor substantiated under Sec. 
27.725.

[Amdt. 27-2, 33 FR 963, Jan. 26, 1968]



Sec. 27.475  Tires and shock absorbers.

    Unless otherwise prescribed, for each specified landing condition, 
the tires must be assumed to be in their static position and the shock 
absorbers to be in their most critical position.



Sec. 27.477  Landing gear arrangement.

    Sections 27.235, 27.479 through 27.485, and 27.493 apply to landing 
gear with two wheels aft, and one or more wheels forward, of the center 
of gravity.



Sec. 27.479  Level landing conditions.

    (a) Attitudes. Under each of the loading conditions prescribed in 
paragraph (b) of this section, the rotorcraft is assumed to be in each 
of the following level landing attitudes:
    (1) An attitude in which all wheels contact the ground 
simultaneously.
    (2) An attitude in which the aft wheels contact the ground with the 
forward wheels just clear of the ground.
    (b) Loading conditions. The rotorcraft must be designed for the 
following landing loading conditions:
    (1) Vertical loads applied under Sec. 27.471.
    (2) The loads resulting from a combination of the loads applied 
under paragraph (b)(1) of this section with drag loads at each wheel of 
not less than 25 percent of the vertical load at that wheel.
    (3) If there are two wheels forward, a distribution of the loads 
applied to those wheels under paragraphs (b)(1) and (2) of this section 
in a ratio of 40:60.
    (c) Pitching moments. Pitching moments are assumed to be resisted 
by--
    (1) In the case of the attitude in paragraph (a)(1) of this section, 
the forward landing gear; and
    (2) In the case of the attitude in paragraph (a)(2) of this section, 
the angular inertia forces.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964; 29 FR 17885, Dec. 17, 1964]



Sec. 27.481  Tail-down landing conditions.

    (a) The rotorcraft is assumed to be in the maximum nose-up attitude 
allowing ground clearance by each part of the rotorcraft.
    (b) In this attitude, ground loads are assumed to act perpendicular 
to the ground.



Sec. 27.483  One-wheel landing conditions.

    For the one-wheel landing condition, the rotorcraft is assumed to be 
in the level attitude and to contact the ground on one aft wheel. In 
this attitude--
    (a) The vertical load must be the same as that obtained on that side 
under Sec. 27.479(b)(1); and
    (b) The unbalanced external loads must be reacted by rotorcraft 
inertia.



Sec. 27.485  Lateral drift landing conditions.

    (a) The rotorcraft is assumed to be in the level landing attitude, 
with--
    (1) Side loads combined with one-half of the maximum ground 
reactions obtained in the level landing conditions of Sec. 27.479 
(b)(1); and
    (2) The loads obtained under paragraph (a)(1) of this section 
applied--
    (i) At the ground contact point; or

[[Page 674]]

    (ii) For full-swiveling gear, at the center of the axle.
    (b) The rotorcraft must be designed to withstand, at ground 
contact--
    (1) When only the aft wheels contact the ground, side loads of 0.8 
times the vertical reaction acting inward on one side, and 0.6 times the 
vertical reaction acting outward on the other side, all combined with 
the vertical loads specified in paragraph (a) of this section; and
    (2) When all wheels contact the ground simultaneously--
    (i) For the aft wheels, the side loads specified in paragraph (b)(1) 
of this section; and
    (ii) For the forward wheels, a side load of 0.8 times the vertical 
reaction combined with the vertical load specified in paragraph (a) of 
this section.



Sec. 27.493  Braked roll conditions.

    Under braked roll conditions with the shock absorbers in their 
static positions--
    (a) The limit vertical load must be based on a load factor of at 
least--
    (1) 1.33, for the attitude specified in Sec. 27.479(a)(1); and
    (2) 1.0 for the attitude specified in Sec. 27.479(a)(2); and
    (b) The structure must be designed to withstand at the ground 
contact point of each wheel with brakes, a drag load at least the lesser 
of--
    (1) The vertical load multiplied by a coefficient of friction of 
0.8; and
    (2) The maximum value based on limiting brake torque.



Sec. 27.497  Ground loading conditions: landing gear with tail wheels.

    (a) General. Rotorcraft with landing gear with two wheels forward, 
and one wheel aft, of the center of gravity must be designed for loading 
conditions as prescribed in this section.
    (b) Level landing attitude with only the forward wheels contacting 
the ground. In this attitude--
    (1) The vertical loads must be applied under Sec. Sec. 27.471 
through 27.475;
    (2) The vertical load at each axle must be combined with a drag load 
at that axle of not less than 25 percent of that vertical load; and
    (3) Unbalanced pitching moments are assumed to be resisted by 
angular inertia forces.
    (c) Level landing attitude with all wheels contacting the ground 
simultaneously. In this attitude, the rotorcraft must be designed for 
landing loading conditions as prescribed in paragraph (b) of this 
section.
    (d) Maximum nose-up attitude with only the rear wheel contacting the 
ground. The attitude for this condition must be the maximum nose-up 
attitude expected in normal operation, including autorotative landings. 
In this attitude--
    (1) The appropriate ground loads specified in paragraphs (b)(1) and 
(2) of this section must be determined and applied, using a rational 
method to account for the moment arm between the rear wheel ground 
reaction and the rotorcraft center of gravity; or
    (2) The probability of landing with initial contact on the rear 
wheel must be shown to be extremely remote.
    (e) Level landing attitude with only one forward wheel contacting 
the ground. In this attitude, the rotorcraft must be designed for ground 
loads as specified in paragraphs (b)(1) and (3) of this section.
    (f) Side loads in the level landing attitude. In the attitudes 
specified in paragraphs (b) and (c) of this section, the following 
apply:
    (1) The side loads must be combined at each wheel with one-half of 
the maximum vertical ground reactions obtained for that wheel under 
paragraphs (b) and (c) of this section. In this condition, the side 
loads must be--
    (i) For the forward wheels, 0.8 times the vertical reaction (on one 
side) acting inward, and 0.6 times the vertical reaction (on the other 
side) acting outward; and
    (ii) For the rear wheel, 0.8 times the vertical reaction.
    (2) The loads specified in paragraph (f)(1) of this section must be 
applied--
    (i) At the ground contact point with the wheel in the trailing 
position (for non-full swiveling landing gear or for full swiveling 
landing gear with a lock, steering device, or shimmy damper to keep the 
wheel in the trailing position); or

[[Page 675]]

    (ii) At the center of the axle (for full swiveling landing gear 
without a lock, steering device, or shimmy damper).
    (g) Braked roll conditions in the level landing attitude. In the 
attitudes specified in paragraphs (b) and (c) of this section, and with 
the shock absorbers in their static positions, the rotorcraft must be 
designed for braked roll loads as follows:
    (1) The limit vertical load must be based on a limit vertical load 
factor of not less than--
    (i) 1.0, for the attitude specified in paragraph (b) of this 
section; and
    (ii) 1.33, for the attitude specified in paragraph (c) of this 
section.
    (2) For each wheel with brakes, a drag load must be applied, at the 
ground contact point, of not less than the lesser of--
    (i) 0.8 times the vertical load; and
    (ii) The maximum based on limiting brake torque.
    (h) Rear wheel turning loads in the static ground attitude. In the 
static ground attitude, and with the shock absorbers and tires in their 
static positions, the rotorcraft must be designed for rear wheel turning 
loads as follows:
    (1) A vertical ground reaction equal to the static load on the rear 
wheel must be combined with an equal sideload.
    (2) The load specified in paragraph (h)(1) of this section must be 
applied to the rear landing gear--
    (i) Through the axle, if there is a swivel (the rear wheel being 
assumed to be swiveled 90 degrees to the longitudinal axis of the 
rotorcraft); or
    (ii) At the ground contact point, if there is a lock, steering 
device or shimmy damper (the rear wheel being assumed to be in the 
trailing position).
    (i) Taxiing condition. The rotorcraft and its landing gear must be 
designed for loads that would occur when the rotorcraft is taxied over 
the roughest ground that may reasonably be expected in normal operation.



Sec. 27.501  Ground loading conditions: landing gear with skids.

    (a) General. Rotorcraft with landing gear with skids must be 
designed for the loading conditions specified in this section. In 
showing compliance with this section, the following apply:
    (1) The design maximum weight, center of gravity, and load factor 
must be determined under Sec. Sec. 27.471 through 27.475.
    (2) Structural yielding of elastic spring members under limit loads 
is acceptable.
    (3) Design ultimate loads for elastic spring members need not exceed 
those obtained in a drop test of the gear with--
    (i) A drop height of 1.5 times that specified in Sec. 27.725; and
    (ii) An assumed rotor lift of not more than 1.5 times that used in 
the limit drop tests prescribed in Sec. 27.725.
    (4) Compliance with paragraphs (b) through (e) of this section must 
be shown with--
    (i) The gear in its most critically deflected position for the 
landing condition being considered; and
    (ii) The ground reactions rationally distributed along the bottom of 
the skid tube.
    (b) Vertical reactions in the level landing attitude. In the level 
attitude, and with the rotorcraft contacting the ground along the bottom 
of both skids, the vertical reactions must be applied as prescribed in 
paragraph (a) of this section.
    (c) Drag reactions in the level landing attitude. In the level 
attitude, and with the rotorcraft contacting the ground along the bottom 
of both skids, the following apply:
    (1) The vertical reactions must be combined with horizontal drag 
reactions of 50 percent of the vertical reaction applied at the ground.
    (2) The resultant ground loads must equal the vertical load 
specified in paragraph (b) of this section.
    (d) Sideloads in the level landing attitude. In the level 
attitude,and with the rotorcraft contacting the ground along the bottom 
of both skids, the following apply:
    (1) The vertical ground reaction must be--
    (i) Equal to the vertical loads obtained in the condition specified 
in paragraph (b) of this section; and
    (ii) Divided equally among the skids.
    (2) The vertical ground reactions must be combined with a horizontal 
sideload of 25 percent of their value.

[[Page 676]]

    (3) The total sideload must be applied equally between the skids and 
along the length of the skids.
    (4) The unbalanced moments are assumed to be resisted by angular 
inertia.
    (5) The skid gear must be investigated for--
    (i) Inward acting sideloads; and
    (ii) Outward acting sideloads.
    (e) One-skid landing loads in the level attitude. In the level 
attitude, and with the rotorcraft contacting the ground along the bottom 
of one skid only, the following apply:
    (1) The vertical load on the ground contact side must be the same as 
that obtained on that side in the condition specified in paragraph (b) 
of this section.
    (2) The unbalanced moments are assumed to be resisted by angular 
inertia.
    (f) Special conditions. In addition to the conditions specified in 
paragraphs (b) and (c) of this section, the rotorcraft must be designed 
for the following ground reactions:
    (1) A ground reaction load acting up and aft at an angle of 45 
degrees to the longitudinal axis of the rotorcraft. This load must be--
    (i) Equal to 1.33 times the maximum weight;
    (ii) Distributed symmetrically among the skids;
    (iii) Concentrated at the forward end of the straight part of the 
skid tube; and
    (iv) Applied only to the forward end of the skid tube and its 
attachment to the rotorcraft.
    (2) With the rotorcraft in the level landing attitude, a vertical 
ground reaction load equal to one-half of the vertical load determined 
under paragraph (b) of this section. This load must be--
    (i) Applied only to the skid tube and its attachment to the 
rotorcraft; and
    (ii) Distributed equally over 33.3 percent of the length between the 
skid tube attachments and centrally located midway between the skid tube 
attachments.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 963, Jan. 26, 1968; Amdt. 27-26, 55 FR 8000, Mar. 6, 1990]



Sec. 27.505  Ski landing conditions.

    If certification for ski operation is requested, the rotorcraft, 
with skis, must be designed to withstand the following loading 
conditions (where P is the maximum static weight on each ski with the 
rotorcraft at design maximum weight, and n is the limit load factor 
determined under Sec. 27.473(b).
    (a) Up-load conditions in which--
    (1) A vertical load of Pn and a horizontal load of Pn/4 are 
simultaneously applied at the pedestal bearings; and
    (2) A vertical load of 1.33 P is applied at the pedestal bearings.
    (b) A side-load condition in which a side load of 0.35 Pn is applied 
at the pedestal bearings in a horizontal plane perpendicular to the 
centerline of the rotorcraft.
    (c) A torque-load condition in which a torque load of 1.33 P (in 
foot pounds) is applied to the ski about the vertical axis through the 
centerline of the pedestal bearings.

                               Water Loads



Sec. 27.521  Float landing conditions.

    If certification for float operation is requested, the rotorcraft, 
with floats, must be designed to withstand the following loading 
conditions (where the limit load factor is determined under Sec. 
27.473(b) or assumed to be equal to that determined for wheel landing 
gear):
    (a) Up-load conditions in which--
    (1) A load is applied so that, with the rotorcraft in the static 
level attitude, the resultant water reaction passes vertically through 
the center of gravity; and
    (2) The vertical load prescribed in paragraph (a)(1) of this section 
is applied simultaneously with an aft component of 0.25 times the 
vertical component.
    (b) A side-load condition in which--
    (1) A vertical load of 0.75 times the total vertical load specified 
in paragraph (a)(1) of this section is divided equally among the floats; 
and
    (2) For each float, the load share determined under paragraph (b)(1) 
of this section, combined with a total side load of 0.25 times the total 
vertical load specified in paragraph (b)(1) of

[[Page 677]]

this section, is applied to that float only.

                       Main Component Requirements



Sec. 27.547  Main rotor structure.

    (a) Each main rotor assembly (including rotor hubs and blades) must 
be designed as prescribed in this section.
    (b) [Reserved]
    (c) The main rotor structure must be designed to withstand the 
following loads prescribed in Sec. Sec. 27.337 through 27.341:
    (1) Critical flight loads.
    (2) Limit loads occurring under normal conditions of autorotation. 
For this condition, the rotor r.p.m. must be selected to include the 
effects of altitude.
    (d) The main rotor structure must be designed to withstand loads 
simulating--
    (1) For the rotor blades, hubs, and flapping hinges, the impact 
force of each blade against its stop during ground operation; and
    (2) Any other critical condition expected in normal operation.
    (e) The main rotor structure must be designed to withstand the limit 
torque at any rotational speed, including zero. In addition:
    (1) The limit torque need not be greater than the torque defined by 
a torque limiting device (where provided), and may not be less than the 
greater of--
    (i) The maximum torque likely to be transmitted to the rotor 
structure in either direction; and
    (ii) The limit engine torque specified in Sec. 27.361.
    (2) The limit torque must be distributed to the rotor blades in a 
rational manner.

(Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, 1425)

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-3, 33 
FR 14105, Sept. 18, 1968]



Sec. 27.549  Fuselage, landing gear, and rotor pylon structures.

    (a) Each fuselage, landing gear, and rotor pylon structure must be 
designed as prescribed in this section. Resultant rotor forces may be 
represented as a single force applied at the rotor hub attachment point.
    (b) Each structure must be designed to withstand--
    (1) The critical loads prescribed in Sec. Sec. 27.337 through 
27.341;
    (2) The applicable ground loads prescribed in Sec. Sec. 27.235, 
27.471 through 27.485, 27.493, 27.497, 27.501, 27.505, and 27.521; and
    (3) The loads prescribed in Sec. 27.547 (d)(2) and (e).
    (c) Auxiliary rotor thrust, and the balancing air and inertia loads 
occurring under accelerated flight conditions, must be considered.
    (d) Each engine mount and adjacent fuselage structure must be 
designed to withstand the loads occurring under accelerated flight and 
landing conditions, including engine torque.

(Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, 1425)

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-3, 33 
FR 14105, Sept. 18, 1968]

                      Emergency Landing Conditions



Sec. 27.561  General.

    (a) The rotorcraft, although it may be damaged in emergency landing 
conditions on land or water, must be designed as prescribed in this 
section to protect the occupants under those conditions.
    (b) The structure must be designed to give each occupant every 
reasonable chance of escaping serious injury in a crash landing when--
    (1) Proper use is made of seats, belts, and other safety design 
provisions;
    (2) The wheels are retracted (where applicable); and
    (3) Each occupant and each item of mass inside the cabin that could 
injure an occupant is restrained when subjected to the following 
ultimate inertial load factors relative to the surrounding structure:
    (i) Upward--4g.
    (ii) Forward--16g.
    (iii) Sideward--8g.
    (iv) Downward--20g, after intended displacement of the seat device.
    (v) Rearward--1.5g.
    (c) The supporting structure must be designed to restrain, under any 
ultimate inertial load up to those specified

[[Page 678]]

in this paragraph, any item of mass above and/or behind the crew and 
passenger compartment that could injure an occupant if it came loose in 
an emergency landing. Items of mass to be considered include, but are 
not limited to, rotors, transmissions, and engines. The items of mass 
must be restrained for the following ultimate inertial load factors:
    (1) Upward--1.5g.
    (2) Forward--12g.
    (3) Sideward--6g.
    (4) Downward--12g.
    (5) Rearward--1.5g
    (d) Any fuselage structure in the area of internal fuel tanks below 
the passenger floor level must be designed to resist the following 
ultimate inertial factors and loads and to protect the fuel tanks from 
rupture when those loads are applied to that area:
    (i) Upward--1.5g.
    (ii) Forward--4.0g.
    (iii) Sideward--2.0g.
    (iv) Downward--4.0g.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-25, 
54 FR 47318, Nov. 13, 1989; Amdt. 27-30, 59 FR 50386, Oct. 3, 1994; 
Amdt. 27-32, 61 FR 10438, Mar. 13, 1996]



Sec. 27.562  Emergency landing dynamic conditions.

    (a) The rotorcraft, although it may be damaged in an emergency crash 
landing, must be designed to reasonably protect each occupant when--
    (1) The occupant properly uses the seats, safety belts, and shoulder 
harnesses provided in the design; and
    (2) The occupant is exposed to the loads resulting from the 
conditions prescribed in this section.
    (b) Each seat type design or other seating device approved for crew 
or passenger occupancy during takeoff and landing must successfully 
complete dynamic tests or be demonstrated by rational analysis based on 
dynamic tests of a similar type seat in accordance with the following 
criteria. The tests must be conducted with an occupant, simulated by a 
170-pound anthropomorphic test dummy (ATD), as defined by 49 CFR 572, 
subpart B, or its equivalent, sitting in the normal upright position.
    (1) A change in downward velocity of not less than 30 feet per 
second when the seat or other seating device is oriented in its nominal 
position with respect to the rotorcraft's reference system, the 
rotorcraft's longitudinal axis is canted upward 60[deg] with respect to 
the impact velocity vector, and the rotorcraft's lateral axis is 
perpendicular to a vertical plane containing the impact velocity vector 
and the rotorcraft's longitudinal axis. Peak floor deceleration must 
occur in not more than 0.031 seconds after impact and must reach a 
minimum of 30g's.
    (2) A change in forward velocity of not less than 42 feet per second 
when the seat or other seating device is oriented in its nominal 
position with respect to the rotorcraft's reference system, the 
rotorcraft's longitudinal axis is yawed 10[deg] either right or left of 
the impact velocity vector (whichever would cause the greatest load on 
the shoulder harness), the rotorcraft's lateral axis is contained in a 
horizontal plane containing the impact velocity vector, and the 
rotorcraft's vertical axis is perpendicular to a horizontal plane 
containing the impact velocity vector. Peak floor deceleration must 
occur in not more than 0.071 seconds after impact and must reach a 
minimum of 18.4g's.
    (3) Where floor rails or floor or sidewall attachment devices are 
used to attach the seating devices to the airframe structure for the 
conditions of this section, the rails or devices must be misaligned with 
respect to each other by at least 10[deg] vertically (i.e., pitch out of 
parallel) and by at least a 10[deg] lateral roll, with the directions 
optional, to account for possible floor warp.
    (c) Compliance with the following must be shown:
    (1) The seating device system must remain intact although it may 
experience separation intended as part of its design.
    (2) The attachment between the seating device and the airframe 
structure must remain intact, although the structure may have exceeded 
its limit load.
    (3) The ATD's shoulder harness strap or straps must remain on or in 
the immediate vicinity of the ATD's shoulder during the impact.

[[Page 679]]

    (4) The safety belt must remain on the ATD's pelvis during the 
impact.
    (5) The ATD's head either does not contact any portion of the crew 
or passenger compartment, or if contact is made, the head impact does 
not exceed a head injury criteria (HIC) of 1,000 as determined by this 
equation.
[GRAPHIC] [TIFF OMITTED] TC28SE91.084

Where: a(t) is the resultant acceleration at the center of gravity of 
          the head form expressed as a multiple of g (the acceleration 
          of gravity) and t2 - t1 is the time 
          duration, in seconds, of major head impact, not to exceed 0.05 
          seconds.

    (6) Loads in individual upper torso harness straps must not exceed 
1,750 pounds. If dual straps are used for retaining the upper torso, the 
total harness strap loads must not exceed 2,000 pounds.
    (7) The maximum compressive load measured between the pelvis and the 
lumbar column of the ATD must not exceed 1,500 pounds.
    (d) An alternate approach that achieves an equivalent or greater 
level of occupant protection, as required by this section, must be 
substantiated on a rational basis.

[Amdt. 27-25, 54 FR 47318, Nov. 13, 1989]



Sec. 27.563  Structural ditching provisions.

    If certification with ditching provisions is requested, structural 
strength for ditching must meet the requirements of this section and 
Sec. 27.801(e).
    (a) Forward speed landing conditions. The rotorcraft must initially 
contact the most critical wave for reasonably probable water conditions 
at forward velocities from zero up to 30 knots in likely pitch, roll, 
and yaw attitudes. The rotorcraft limit vertical descent velocity may 
not be less than 5 feet per second relative to the mean water surface. 
Rotor lift may be used to act through the center of gravity throughout 
the landing impact. This lift may not exceed two-thirds of the design 
maximum weight. A maximum forward velocity of less than 30 knots may be 
used in design if it can be demonstrated that the forward velocity 
selected would not be exceeded in a normal one-engine-out touchdown.
    (b) Auxiliary or emergency float conditions--(1) Floats fixed or 
deployed before initial water contact. In addition to the landing loads 
in paragraph (a) of this section, each auxiliary or emergency float, of 
its support and attaching structure in the airframe or fuselage, must be 
designed for the load developed by a fully immersed float unless it can 
be shown that full immersion is unlikely. If full immersion is unlikely, 
the highest likely float buoyancy load must be applied. The highest 
likely buoyancy load must include consideration of a partially immersed 
float creating restoring moments to compensate the upsetting moments 
caused by side wind, unsymmetrical rotorcraft loading, water wave 
action, rotorcraft inertia, and probable structural damage and leakage 
considered under Sec. 27.801(d). Maximum roll and pitch angles 
determined from compliance with Sec. 27.801(d) may be used, if 
significant, to determine the extent of immersion of each float. If the 
floats are deployed in flight, appropriate air loads derived from the 
flight limitations with the floats deployed shall be used in 
substantiation of the floats and their attachment to the rotorcraft. For 
this purpose, the design airspeed for limit load is the float deployed 
airspeed operating limit multiplied by 1.11.
    (2) Floats deployed after initial water contact. Each float must be 
designed for full or partial immersion perscribed in paragraph (b)(1) of 
this section. In addition, each float must be designed for combined 
vertical and drag loads using a relative limit speed of 20 knots between 
the rotorcraft and the water. The vertical load may not be less than the 
highest likely buoyancy load determined under paragraph (b)(1) of this 
section.

[Amdt. 27-26, 55 FR 8000, Mar. 6, 1990]

                           Fatigue Evaluation



Sec. 27.571  Fatigue evaluation of flight structure.

    (a) General. Each portion of the flight structure (the flight 
structure includes rotors, rotor drive systems between the engines and 
the rotor hubs, controls,

[[Page 680]]

fuselage, landing gear, and their related primary attachments), the 
failure of which could be catastrophic, must be identified and must be 
evaluated under paragraph (b), (c), (d), or (e) of this section. The 
following apply to each fatigue evaluation:
    (1) The procedure for the evaluation must be approved.
    (2) The locations of probable failure must be determined.
    (3) Inflight measurement must be included in determining the 
following:
    (i) Loads or stresses in all critical conditions throughout the 
range of limitations in Sec. 27.309, except that maneuvering load 
factors need not exceed the maximum values expected in operation.
    (ii) The effect of altitude upon these loads or stresses.
    (4) The loading spectra must be as severe as those expected in 
operation including, but not limited to, external cargo operations, if 
applicable, and ground-air-ground cycles. The loading spectra must be 
based on loads or stresses determined under paragraph (a)(3) of this 
section.
    (b) Fatigue tolerance evaluation. It must be shown that the fatigue 
tolerance of the structure ensures that the probability of catastrophic 
fatigue failure is extremely remote without establishing replacement 
times, inspection intervals or other procedures under section A27.4 of 
appendix A.
    (c) Replacement time evaluation. it must be shown that the 
probability of catastrophic fatigue failure is extremely remote within a 
replacement time furnished under section A27.4 of appendix A.
    (d) Fail-safe evaluation. The following apply to fail-safe 
evaluation:
    (1) It must be shown that all partial failures will become readily 
detectable under inspection procedures furnished under section A27.4 of 
appendix A.
    (2) The interval between the time when any partial failure becomes 
readily detectable under paragraph (d)(1) of this section, and the time 
when any such failure is expected to reduce the remaining strength of 
the structure to limit or maximum attainable loads (whichever is less), 
must be determined.
    (3) It must be shown that the interval determined under paragraph 
(d)(2) of this section is long enough, in relation to the inspection 
intervals and related procedures furnished under section A27.4 of 
appendix A, to provide a probability of detection great enough to ensure 
that the probability of catastrophic failure is extremely remote.
    (e) Combination of replacement time and failsafe evaluations. A 
component may be evaluated under a combination of paragraphs (c) and (d) 
of this section. For such component it must be shown that the 
probability of catastrophic failure is extremely remote with an approved 
combination of replacement time, inspection intervals, and related 
procedures furnished under section A27.4 of appendix A.

(Secs. 313(a), 601, 603, 604, and 605, 72 Stat. 752, 775, and 778, (49 
U.S.C. 1354(a), 1421, 1423, 1424, and 1425; sec. 6(c), 49 U.S.C. 
1655(c)))

[Amdt. 27-3, 33 FR 14106, Sept. 18, 1968, as amended by Amdt. 27-12, 42 
FR 15044, Mar. 17, 1977; Amdt. 27-18, 45 FR 60177, Sept. 11 1980; Amdt. 
27-26, 55 FR 8000, Mar. 6, 1990]



Sec. 27.573  Damage Tolerance and Fatigue Evaluation of Composite 
Rotorcraft Structures.

    (a) Each applicant must evaluate the composite rotorcraft structure 
under the damage tolerance standards of paragraph (d) of this section 
unless the applicant establishes that a damage tolerance evaluation is 
impractical within the limits of geometry, inspectability, and good 
design practice. If an applicant establishes that it is impractical 
within the limits of geometry, inspectability, and good design practice, 
the applicant must do a fatigue evaluation in accordance with paragraph 
(e) of this section.
    (b) The methodology used to establish compliance with this section 
must be submitted to and approved by the Administrator.
    (c) Definitions:
    (1) Catastrophic failure is an event that could prevent continued 
safe flight and landing.
    (2) Principal Structural Elements (PSEs) are structural elements 
that contribute significantly to the carrying of flight or ground loads, 
the failure of which could result in catastrophic failure of the 
rotorcraft.

[[Page 681]]

    (3) Threat Assessment is an assessment that specifies the locations, 
types, and sizes of damage, considering fatigue, environmental effects, 
intrinsic and discrete flaws, and impact or other accidental damage 
(including the discrete source of the accidental damage) that may occur 
during manufacture or operation.
    (d) Damage Tolerance Evaluation:
    (1) Each applicant must show that catastrophic failure due to static 
and fatigue loads, considering the intrinsic or discrete manufacturing 
defects or accidental damage, is avoided throughout the operational life 
or prescribed inspection intervals of the rotorcraft by performing 
damage tolerance evaluations of the strength of composite PSEs and other 
parts, detail design points, and fabrication techniques. Each applicant 
must account for the effects of material and process variability along 
with environmental conditions in the strength and fatigue evaluations. 
Each applicant must evaluate parts that include PSEs of the airframe, 
main and tail rotor drive systems, main and tail rotor blades and hubs, 
rotor controls, fixed and movable control surfaces, engine and 
transmission mountings, landing gear, other parts, detail design points, 
and fabrication techniques deemed critical by the FAA. Each damage 
tolerance evaluation must include:
    (i) The identification of all PSEs;
    (ii) In-flight and ground measurements for determining the loads or 
stresses for all PSEs for all critical conditions throughout the range 
of limits in Sec. 27.309 (including altitude effects), except that 
maneuvering load factors need not exceed the maximum values expected in 
service;
    (iii) The loading spectra as severe as those expected in service 
based on loads or stresses determined under paragraph (d)(1)(ii) of this 
section, including external load operations, if applicable, and other 
operations including high-torque events;
    (iv) A threat assessment for all PSEs that specifies the locations, 
types, and sizes of damage, considering fatigue, environmental effects, 
intrinsic and discrete flaws, and impact or other accidental damage 
(including the discrete source of the accidental damage) that may occur 
during manufacture or operation; and
    (v) An assessment of the residual strength and fatigue 
characteristics of all PSEs that supports the replacement times and 
inspection intervals established under paragraph (d)(2) of this section.
    (2) Each applicant must establish replacement times, inspections, or 
other procedures for all PSEs to require the repair or replacement of 
damaged parts before a catastrophic failure. These replacement times, 
inspections, or other procedures must be included in the Airworthiness 
Limitations Section of the Instructions for Continued Airworthiness 
required by Sec. 27.1529.
    (i) Replacement times for PSEs must be determined by tests, or by 
analysis supported by tests, and must show that the structure is able to 
withstand the repeated loads of variable magnitude expected in-service. 
In establishing these replacement times, the following items must be 
considered:
    (A) Damage identified in the threat assessment required by paragraph 
(d)(1)(iv) of this section;
    (B) Maximum acceptable manufacturing defects and in-service damage 
(i.e., those that do not lower the residual strength below ultimate 
design loads and those that can be repaired to restore ultimate 
strength); and
    (C) Ultimate load strength capability after applying repeated loads.
    (ii) Inspection intervals for PSEs must be established to reveal any 
damage identified in the threat assessment required by paragraph 
(d)(1)(iv) of this section that may occur from fatigue or other in-
service causes before such damage has grown to the extent that the 
component cannot sustain the required residual strength capability. In 
establishing these inspection intervals, the following items must be 
considered:
    (A) The growth rate, including no-growth, of the damage under the 
repeated loads expected in-service determined by tests or analysis 
supported by tests;
    (B) The required residual strength for the assumed damage 
established after considering the damage type, inspection interval, 
detectability of damage, and the techniques adopted for damage

[[Page 682]]

detection. The minimum required residual strength is limit load; and
    (C) Whether the inspection will detect the damage growth before the 
minimum residual strength is reached and restored to ultimate load 
capability, or whether the component will require replacement.
    (3) Each applicant must consider the effects of damage on stiffness, 
dynamic behavior, loads, and functional performance on all PSEs when 
substantiating the maximum assumed damage size and inspection interval.
    (e) Fatigue Evaluation: If an applicant establishes that the damage 
tolerance evaluation described in paragraph (d) of this section is 
impractical within the limits of geometry, inspectability, or good 
design practice, the applicant must do a fatigue evaluation of the 
particular composite rotorcraft structure and:
    (1) Identify all PSEs considered in the fatigue evaluation;
    (2) Identify the types of damage for all PSEs considered in the 
fatigue evaluation;
    (3) Establish supplemental procedures to minimize the risk of 
catastrophic failure associated with the damages identified in paragraph 
(d) of this section; and
    (4) Include these supplemental procedures in the Airworthiness 
Limitations section of the Instructions for Continued Airworthiness 
required by Sec. 27.1529.

[Doc. No. FAA-2009-0660, Amdt. 27-47, 76 FR 74663, Dec. 1, 2011]



                    Subpart D_Design and Construction

                                 General



Sec. 27.601  Design.

    (a) The rotorcraft may have no design features or details that 
experience has shown to be hazardous or unreliable.
    (b) The suitability of each questionable design detail and part must 
be established by tests.



Sec. 27.602  Critical parts.

    (a) Critical part. A critical part is a part, the failure of which 
could have a catastrophic effect upon the rotocraft, and for which 
critical characteristics have been identified which must be controlled 
to ensure the required level of integrity.
    (b) If the type design includes crtical parts, a critical parts list 
shall be established. Procedures shall be established to define the 
critical design characteristics, identify processes that affect those 
characteristics, and identify the design change and process change 
controls necessary for showing compliance with the quality assurance 
requirements of part 21 of this chapter.

[Doc. No. 29311, 64 FR 46232, Aug. 24, 1999]



Sec. 27.603  Materials.

    The suitability and durability of materials used for parts, the 
failure of which could adversely affect safety, must--
    (a) Be established on the basis of experience or tests;
    (b) Meet approved specifications that ensure their having the 
strength and other properties assumed in the design data; and
    (c) Take into account the effects of environmental conditions, such 
as temperature and humidity, expected in service.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424); and sec. 6(c) of the Dept. of Transportation 
Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55469, Dec. 20, 1976; Amdt. 27-16, 43 FR 50599, Oct. 30, 1978]



Sec. 27.605  Fabrication methods.

    (a) The methods of fabrication used must produce consistently sound 
structures. If a fabrication process (such as gluing, spot welding, or 
heat-treating) requires close control to reach this objective, the 
process must be performed according to an approved process 
specification.
    (b) Each new aircraft fabrication method must be substantiated by a 
test program.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424 and 1425); sec. 6(c) of the 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 
43 FR 50599, Oct. 30, 1978]

[[Page 683]]



Sec. 27.607  Fasteners.

    (a) Each removable bolt, screw, nut, pin, or other fastener whose 
loss could jeopardize the safe operation of the rotorcraft must 
incorporate two separate locking devices. The fastener and its locking 
devices may not be adversely affected by the environmental conditions 
associated with the particular installation.
    (b) No self-locking nut may be used on any bolt subject to rotation 
in operation unless a nonfriction locking device is used in addition to 
the self-locking device.

[Amdt. 27-4, 33 FR 14533, Sept. 27, 1968]



Sec. 27.609  Protection of structure.

    Each part of the structure must--
    (a) Be suitably protected against deterioration or loss of strength 
in service due to any cause, including--
    (1) Weathering;
    (2) Corrosion; and
    (3) Abrasion; and
    (b) Have provisions for ventilation and drainage where necessary to 
prevent the accumulation of corrosive, flammable, or noxious fluids.



Sec. 27.610  Lightning and static electricity protection.

    (a) The rotorcraft must be protected against catastrophic effects 
from lightning.
    (b) For metallic components, compliance with paragraph (a) of this 
section may be shown by--
    (1) Electrically bonding the components properly to the airframe; or
    (2) Designing the components so that a strike will not endanger the 
rotorcraft.
    (c) For nonmetallic components, compliance with paragraph (a) of 
this section may be shown by--
    (1) Designing the components to minimize the effect of a strike; or
    (2) Incorporating acceptable means of diverting the resulting 
electrical current so as not to endanger the rotorcraft.
    (d) The electrical bonding and protection against lightning and 
static electricity must--
    (1) Minimize the accumulation of electrostatic charge;
    (2) Minimize the risk of electric shock to crew, passengers, and 
service and maintenance personnel using normal precautions;
    (3) Provide an electrical return path, under both normal and fault 
conditions, on rotorcraft having grounded electrical systems; and
    (4) Reduce to an acceptable level the effects of static electricity 
on the functioning of essential electrical and electronic equipment.

[Amdt. 27-21, 49 FR 44433, Nov. 6, 1984, as amended by Amdt. 27-37, 64 
FR 45094, Aug. 18, 1999; Amdt. 27-46, 76 FR 33135, June 8, 2011]



Sec. 27.611  Inspection provisions.

    There must be means to allow the close examination of each part that 
requires--
    (a) Recurring inspection;
    (b) Adjustment for proper alignment and functioning; or
    (c) Lubrication.



Sec. 27.613  Material strength properties and design values.

    (a) Material strength properties must be based on enough tests of 
material meeting specifications to establish design values on a 
statistical basis.
    (b) Design values must be chosen to minimize the probability of 
structural failure due to material variability. Except as provided in 
paragraphs (d) and (e) of this section, compliance with this paragraph 
must be shown by selecting design values that assure material strength 
with the following probability--
    (1) Where applied loads are eventually distributed through a single 
member within an assembly, the failure of which would result in loss of 
structural integrity of the component, 99 percent probability with 95 
percent confidence; and
    (2) For redundant structure, those in which the failure of 
individual elements would result in applied loads being safely 
distributed to other load-carrying members, 90 percent probability with 
95 percent confidence.
    (c) The strength, detail design, and fabrication of the structure 
must minimize the probability of disastrous fatigue failure, 
particularly at points of stress concentration.

[[Page 684]]

    (d) Design values may be those contained in the following 
publications (available from the Naval Publications and Forms Center, 
5801 Tabor Avenue, Philadelphia, Pennsylvania 19120) or other values 
approved by the Administrator:
    (1) MIL-HDBK-5, ``Metallic Materials and Elements for Flight Vehicle 
Structure''.
    (2) MIL-HDBK-17, ``Plastics for Flight Vehicles''.
    (3) ANC-18, ``Design of Wood Aircraft Structures''.
    (4) MIL-HDBK-23, ``Composite Construction for Flight Vehicles''.
    (e) Other design values may be used if a selection of the material 
is made in which a specimen of each individual item is tested before use 
and it is determined that the actual strength properties of that 
particular item will equal or exceed those used in design.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 
43 FR 50599, Oct. 30, 1978; Amdt. 27-26, 55 FR 8000, Mar. 6, 1990]



Sec. 27.619  Special factors.

    (a) The special factors prescribed in Sec. Sec. 27.621 through 
27.625 apply to each part of the structure whose strength is--
    (1) Uncertain;
    (2) Likely to deteriorate in service before normal replacement; or
    (3) Subject to appreciable variability due to--
    (i) Uncertainties in manufacturing processes; or
    (ii) Uncertainties in inspection methods.
    (b) For each part to which Sec. Sec. 27.621 through 27.625 apply, 
the factor of safety prescribed in Sec. 27.303 must be multiplied by a 
special factor equal to--
    (1) The applicable special factors prescribed in Sec. Sec. 27.621 
through 27.625; or
    (2) Any other factor great enough to ensure that the probability of 
the part being understrength because of the uncertainties specified in 
paragraph (a) of this section is extremely remote.



Sec. 27.621  Casting factors.

    (a) General. The factors, tests, and inspections specified in 
paragraphs (b) and (c) of this section must be applied in addition to 
those necessary to establish foundry quality control. The inspections 
must meet approved specifications. Paragraphs (c) and (d) of this 
section apply to structural castings except castings that are pressure 
tested as parts of hydraulic or other fluid systems and do not support 
structural loads.
    (b) Bearing stresses and surfaces. The casting factors specified in 
paragraphs (c) and (d) of this section--
    (1) Need not exceed 1.25 with respect to bearing stresses regardless 
of the method of inspection used; and
    (2) Need not be used with respect to the bearing surfaces of a part 
whose bearing factor is larger than the applicable casting factor.
    (c) Critical castings. For each casting whose failure would preclude 
continued safe flight and landing of the rotorcraft or result in serious 
injury to any occupant, the following apply:
    (1) Each critical casting must--
    (i) Have a casting factor of not less than 1.25; and
    (ii) Receive 100 percent inspection by visual, radiographic, and 
magnetic particle (for ferromagnetic materials) or penetrant (for 
nonferromagnetic materials) inspection methods or approved equivalent 
inspection methods.
    (2) For each critical casting with a casting factor less than 1.50, 
three sample castings must be static tested and shown to meet--
    (i) The strength requirements of Sec. 27.305 at an ultimate load 
corresponding to a casting factor of 1.25; and
    (ii) The deformation requirements of Sec. 27.305 at a load of 1.15 
times the limit load.
    (d) Noncritical castings. For each casting other than those 
specified in paragraph (c) of this section, the following apply:
    (1) Except as provided in paragraphs (d)(2) and (3) of this section, 
the casting factors and corresponding inspections must meet the 
following table:

------------------------------------------------------------------------
              Casting factor                         Inspection
------------------------------------------------------------------------
2.0 or greater...........................  100 percent visual.

[[Page 685]]

 
Less than 2.0, greater than 1.5..........  100 percent visual, and
                                            magnetic particle
                                            (ferromagnetic materials),
                                            penetrant (nonferromagnetic
                                            materials), or approved
                                            equivalent inspection
                                            methods.
1.25 through 1.50........................  100 percent visual, and
                                            magnetic particle
                                            (ferromagnetic materials).
                                            penetrant (nonferromagnetic
                                            materials), and radiographic
                                            or approved equivalent
                                            inspection methods.
------------------------------------------------------------------------

    (2) The percentage of castings inspected by nonvisual methods may be 
reduced below that specified in paragraph (d)(1) of this section when an 
approved quality control procedure is established.
    (3) For castings procured to a specification that guarantees the 
mechanical properties of the material in the casting and provides for 
demonstration of these properties by test of coupons cut from the 
castings on a sampling basis--
    (i) A casting factor of 1.0 may be used; and
    (ii) The castings must be inspected as provided in paragraph (d)(1) 
of this section for casting factors of ``1.25 through 1.50'' and tested 
under paragraph (c)(2) of this section.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-34, 
62 FR 46173, Aug. 29, 1997]



Sec. 27.623  Bearing factors.

    (a) Except as provided in paragraph (b) of this section, each part 
that has clearance (free fit), and that is subject to pounding or 
vibration, must have a bearing factor large enough to provide for the 
effects of normal relative motion.
    (b) No bearing factor need be used on a part for which any larger 
special factor is prescribed.



Sec. 27.625  Fitting factors.

    For each fitting (part or terminal used to join one structural 
member to another) the following apply:
    (a) For each fitting whose strength is not proven by limit and 
ultimate load tests in which actual stress conditions are simulated in 
the fitting and surrounding structures, a fitting factor of at least 
1.15 must be applied to each part of--
    (1) The fitting;
    (2) The means of attachment; and
    (3) The bearing on the joined members.
    (b) No fitting factor need be used--
    (1) For joints made under approved practices and based on 
comprehensive test data (such as continuous joints in metal plating, 
welded joints, and scarf joints in wood); and
    (2) With respect to any bearing surface for which a larger special 
factor is used.
    (c) For each integral fitting, the part must be treated as a fitting 
up to the point at which the section properties become typical of the 
member.
    (d) Each seat, berth, litter, safety belt, and harness attachment to 
the structure must be shown by analysis, tests, or both, to be able to 
withstand the inertia forces prescribed in Sec. 27.561(b)(3) multiplied 
by a fitting factor of 1.33.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-35, 
63 FR 43285, Aug. 12, 1998]



Sec. 27.629  Flutter.

    Each aerodynamic surface of the rotorcraft must be free from flutter 
under each appropriate speed and power condition.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-26, 
55 FR 8000, Mar. 6, 1990]

                                 Rotors



Sec. 27.653  Pressure venting and drainage of rotor blades.

    (a) For each rotor blade--
    (1) There must be means for venting the internal pressure of the 
blade;
    (2) Drainage holes must be provided for the blade; and
    (3) The blade must be designed to prevent water from becoming 
trapped in it.
    (b) Paragraphs (a)(1) and (2) of this section does not apply to 
sealed rotor blades capable of withstanding the maximum pressure 
differentials expected in service.

[Amdt. 27-2, 33 FR 963, Jan. 26, 1968]



Sec. 27.659  Mass balance.

    (a) The rotors and blades must be mass balanced as necessary to--

[[Page 686]]

    (1) Prevent excessive vibration; and
    (2) Prevent flutter at any speed up to the maximum forward speed.
    (b) The structural integrity of the mass balance installation must 
be substantiated.

[Amdt. 27-2, 33 FR 963, Jan. 26, 1968]



Sec. 27.661  Rotor blade clearance.

    There must be enough clearance between the rotor blades and other 
parts of the structure to prevent the blades from striking any part of 
the structure during any operating condition.

[Amdt. 27-2, 33 FR 963, Jan. 26, 1968]



Sec. 27.663  Ground resonance prevention means.

    (a) The reliability of the means for preventing ground resonance 
must be shown either by analysis and tests, or reliable service 
experience, or by showing through analysis or tests that malfunction or 
failure of a single means will not cause ground resonance.
    (b) The probable range of variations, during service, of the damping 
action of the ground resonance prevention means must be established and 
must be investigated during the test required by Sec. 27.241.

[Amdt. 27-2, 33 FR 963, Jan. 26, 1968, as amended by Amdt. 27-26, 55 FR 
8000, Mar. 6, 1990]

                             Control Systems



Sec. 27.671  General.

    (a) Each control and control system must operate with the ease, 
smoothness, and positiveness appropriate to its function.
    (b) Each element of each flight control system must be designed, or 
distinctively and permanently marked, to minimize the probability of any 
incorrect assembly that could result in the malfunction of the system.



Sec. 27.672  Stability augmentation, automatic, and power-operated systems.

    If the functioning of stability augmentation or other automatic or 
power-operated systems is necessary to show compliance with the flight 
characteristics requirements of this part, such systems must comply with 
Sec. 27.671 of this part and the following:
    (a) A warning which is clearly distinguishable to the pilot under 
expected flight conditions without requiring the pilot's attention must 
be provided for any failure in the stability augmentation system or in 
any other automatic or power-operated system which could result in an 
unsafe condition if the pilot is unaware of the failure. Warning systems 
must not activate the control systems.
    (b) The design of the stability augmentation system or of any other 
automatic or power-operated system must allow initial counteraction of 
failures without requiring exceptional pilot skill or strength by 
overriding the failure by movement of the flight controls in the normal 
sense and deactivating the failed system.
    (c) It must be shown that after any single failure of the stability 
augmentation system or any other automatic or power-operated system--
    (1) The rotorcraft is safely controllable when the failure or 
malfunction occurs at any speed or altitude within the approved 
operating limitations;
    (2) The controllability and maneuverability requirements of this 
part are met within a practical operational flight envelope (for 
example, speed, altitude, normal acceleration, and rotorcraft 
configurations) which is described in the Rotorcraft Flight Manual; and
    (3) The trim and stability characteristics are not impaired below a 
level needed to permit continued safe flight and landing.

[Amdt. 27-21, 49 FR 44433, Nov. 6, 1984; 49 FR 47594, Dec. 6, 1984]



Sec. 27.673  Primary flight control.

    Primary flight controls are those used by the pilot for immediate 
control of pitch, roll, yaw, and vertical motion of the rotorcraft.

[Amdt. 27-21, 49 FR 44434, Nov. 6, 1984]



Sec. 27.674  Interconnected controls.

    Each primary flight control system must provide for safe flight and 
landing and operate independently after a malfunction, failure, or jam 
of any auxiliary interconnected control.

[Amdt. 27-26, 55 FR 8001, Mar. 6, 1990]

[[Page 687]]



Sec. 27.675  Stops.

    (a) Each control system must have stops that positively limit the 
range of motion of the pilot's controls.
    (b) Each stop must be located in the system so that the range of 
travel of its control is not appreciably affected by--
    (1) Wear;
    (2) Slackness; or
    (3) Takeup adjustments.
    (c) Each stop must be able to withstand the loads corresponding to 
the design conditions for the system.
    (d) For each main rotor blade--
    (1) Stops that are appropriate to the blade design must be provided 
to limit travel of the blade about its hinge points; and
    (2) There must be means to keep the blade from hitting the droop 
stops during any operation other than starting and stopping the rotor.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 
43 FR 50599, Oct. 30, 1978]



Sec. 27.679  Control system locks.

    If there is a device to lock the control system with the rotorcraft 
on the ground or water, there must be means to--
    (a) Give unmistakable warning to the pilot when the lock is engaged; 
and
    (b) Prevent the lock from engaging in flight.



Sec. 27.681  Limit load static tests.

    (a) Compliance with the limit load requirements of this part must be 
shown by tests in which--
    (1) The direction of the test loads produces the most severe loading 
in the control system; and
    (2) Each fitting, pulley, and bracket used in attaching the system 
to the main structure is included.
    (b) Compliance must be shown (by analyses or individual load tests) 
with the special factor requirements for control system joints subject 
to angular motion.



Sec. 27.683  Operation tests.

    It must be shown by operation tests that, when the controls are 
operated from the pilot compartment with the control system loaded to 
correspond with loads specified for the system, the system is free 
from--
    (a) Jamming;
    (b) Excessive friction; and
    (c) Excessive deflection.



Sec. 27.685  Control system details.

    (a) Each detail of each control system must be designed to prevent 
jamming, chafing, and interference from cargo, passengers, loose objects 
or the freezing of moisture.
    (b) There must be means in the cockpit to prevent the entry of 
foreign objects into places where they would jam the system.
    (c) There must be means to prevent the slapping of cables or tubes 
against other parts.
    (d) Cable systems must be designed as follows:
    (1) Cables, cable fittings, turnbuckles, splices, and pulleys must 
be of an acceptable kind.
    (2) The design of the cable systems must prevent any hazardous 
change in cable tension throughout the range of travel under any 
operating conditions and temperature variations.
    (3) No cable smaller than three thirty-seconds of an inch diameter 
may be used in any primary control system.
    (4) Pulley kinds and sizes must correspond to the cables with which 
they are used. The pulley cable combinations and strength values which 
must be used are specified in Military Handbook MIL-HDBK-5C, Vol. 1 & 
Vol. 2, Metallic Materials and Elements for Flight Vehicle Structures, 
(Sept. 15, 1976, as amended through December 15, 1978). This 
incorporation by reference was approved by the Director of the Federal 
Register in accordance with 5 U.S.C. section 552(a) and 1 CFR part 51. 
Copies may be obtained from the Naval Publications and Forms Center, 
5801 Tabor Avenue, Philadelphia, Pennsylvania, 19120. Copies may be 
inspected at the FAA, Rotorcraft Standards Staff, 4400 Blue Mount Road, 
Fort Worth, Texas, or at the National Archives and Records 
Administration

[[Page 688]]

(NARA). For information on the availability of this material at NARA, 
call 202-741-6030, or go to: http://www.archives.gov/federal--register/
code--of--federal--regulations/ibr--locations.html.
    (5) Pulleys must have close fitting guards to prevent the cables 
from being displaced or fouled.
    (6) Pulleys must lie close enough to the plane passing through the 
cable to prevent the cable from rubbing against the pulley flange.
    (7) No fairlead may cause a change in cable direction of more than 
3[deg].
    (8) No clevis pin subject to load or motion and retained only by 
cotter pins may be used in the control system.
    (9) Turnbuckles attached to parts having angular motion must be 
installed to prevent binding throughout the range of travel.
    (10) There must be means for visual inspection at each fairlead, 
pulley, terminal, and turnbuckle.
    (e) Control system joints subject to angular motion must incorporate 
the following special factors with respect to the ultimate bearing 
strength of the softest material used as a bearing:
    (1) 3.33 for push-pull systems other than ball and roller bearing 
systems.
    (2) 2.0 for cable systems.
    (f) For control system joints, the manufacturer's static, non-
Brinell rating of ball and roller bearings must not be exceeded.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55469, Dec. 20, 1976; Amdt. 27-26, 55 FR 8001, Mar. 6, 1990; 69 FR 
18803, Apr. 9, 2004]



Sec. 27.687  Spring devices.

    (a) Each control system spring device whose failure could cause 
flutter or other unsafe characteristics must be reliable.
    (b) Compliance with paragraph (a) of this section must be shown by 
tests simulating service conditions.



Sec. 27.691  Autorotation control mechanism.

    Each main rotor blade pitch control mechanism must allow rapid entry 
into autorotation after power failure.



Sec. 27.695  Power boost and power-operated control system.

    (a) If a power boost or power-operated control system is used, an 
alternate system must be immediately available that allows continued 
safe flight and landing in the event of--
    (1) Any single failure in the power portion of the system; or
    (2) The failure of all engines.
    (b) Each alternate system may be a duplicate power portion or a 
manually operated mechanical system. The power portion includes the 
power source (such as hydraulic pumps), and such items as valves, lines, 
and actuators.
    (c) The failure of mechanical parts (such as piston rods and links), 
and the jamming of power cylinders, must be considered unless they are 
extremely improbable.

                              Landing Gear



Sec. 27.723  Shock absorption tests.

    The landing inertia load factor and the reserve energy absorption 
capacity of the landing gear must be substantiated by the tests 
prescribed in Sec. Sec. 27.725 and 27.727, respectively. These tests 
must be conducted on the complete rotorcraft or on units consisting of 
wheel, tire, and shock absorber in their proper relation.



Sec. 27.725  Limit drop test.

    The limit drop test must be conducted as follows:
    (a) The drop height must be--
    (1) 13 inches from the lowest point of the landing gear to the 
ground; or
    (2) Any lesser height, not less than eight inches, resulting in a 
drop contact velocity equal to the greatest probable sinking speed 
likely to occur at ground contact in normal power-off landings.
    (b) If considered, the rotor lift specified in Sec. 27.473(a) must 
be introduced into the drop test by appropriate energy absorbing devices 
or by the use of an effective mass.
    (c) Each landing gear unit must be tested in the attitude simulating 
the landing condition that is most critical from the standpoint of the 
energy to be absorbed by it.

[[Page 689]]

    (d) When an effective mass is used in showing compliance with 
paragraph (b) of this section, the following formula may be used instead 
of more rational computations:
[GRAPHIC] [TIFF OMITTED] TC28SE91.085

where:

We=the effective weight to be used in the drop test (lbs.);
W=WM for main gear units (lbs.), equal to the static reaction on the 
          particular unit with the rotorcraft in the most critical 
          attitude. A rational method may be used in computing a main 
          gear static reaction, taking into consideration the moment arm 
          between the main wheel reaction and the rotorcraft center of 
          gravity.
W=WN for nose gear units (lbs.), equal to the vertical component of the 
          static reaction that would exist at the nose wheel, assuming 
          that the mass of the rotorcraft acts at the center of gravity 
          and exerts a force of 1.0g downward and 0.25g forward.
W=WT for tailwheel units (lbs.), equal to whichever of the following is 
          critical:

    (1) The static weight on the tailwheel with the rotorcraft resting 
on all wheels; or
    (2) The vertical component of the ground reaction that would occur 
at the tailwheel, assuming that the mass of the rotorcraft acts at the 
center of gravity and exerts a force of lg downward with the rotorcraft 
in the maximum nose-up attitude considered in the nose-up landing 
conditions.

h=specified free drop height (inches).
L=ration of assumed rotor lift to the rotorcraft weight.
d=deflection under impact of the tire (at the proper inflation pressure) 
          plus the vertical component of the axle travels (inches) 
          relative to the drop mass.
n=limit inertia load factor.
nj=the load factor developed, during impact, on the mass used in the 
          drop test (i.e., the acceleration dv/dt in g's recorded in the 
          drop test plus 1.0).



Sec. 27.727  Reserve energy absorption drop test.

    The reserve energy absorption drop test must be conducted as 
follows:
    (a) The drop height must be 1.5 times that specified in Sec. 
27.725(a).
    (b) Rotor lift, where considered in a manner similar to that 
prescribed in Sec. 27.725(b), may not exceed 1.5 times the lift allowed 
under that paragraph.
    (c) The landing gear must withstand this test without collapsing. 
Collapse of the landing gear occurs when a member of the nose, tail, or 
main gear will not support the rotorcraft in the proper attitude or 
allows the rotorcraft structure, other than the landing gear and 
external accessories, to impact the landing surface.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-26, 
55 FR 8001, Mar. 6, 1990]



Sec. 27.729  Retracting mechanism.

    For rotorcraft with retractable landing gear, the following apply:
    (a) Loads. The landing gear, retracting mechansim, wheel-well doors, 
and supporting structure must be designed for--
    (1) The loads occurring in any maneuvering condition with the gear 
retracted;
    (2) The combined friction, inertia, and air loads occurring during 
retraction and extension at any airspeed up to the design maximum 
landing gear operating speed; and
    (3) The flight loads, including those in yawed flight, occurring 
with the gear extended at any airspeed up to the design maximum landing 
gear extended speed.
    (b) Landing gear lock. A positive means must be provided to keep the 
gear extended.
    (c) Emergency operation. When other than manual power is used to 
operate the gear, emergency means must be provided for extending the 
gear in the event of--
    (1) Any reasonably probable failure in the normal retraction system; 
or
    (2) The failure of any single source of hydraulic, electric, or 
equivalent energy.
    (d) Operation tests. The proper functioning of the retracting 
mechanism must be shown by operation tests.
    (e) Position indicator. There must be a means to indicate to the 
pilot when the gear is secured in the extreme positions.
    (f) Control. The location and operation of the retraction control 
must meet the requirements of Sec. Sec. 27.777 and 27.779.
    (g) Landing gear warning. An aural or equally effective landing gear 
warning

[[Page 690]]

device must be provided that functions continuously when the rotorcraft 
is in a normal landing mode and the landing gear is not fully extended 
and locked. A manual shutoff capability must be provided for the warning 
device and the warning system must automatically reset when the 
rotorcraft is no longer in the landing mode.

[Amdt. 27-21, 49 FR 44434, Nov. 6, 1984]



Sec. 27.731  Wheels.

    (a) Each landing gear wheel must be approved.
    (b) The maximum static load rating of each wheel may not be less 
than the corresponding static ground reaction with--
    (1) Maximum weight; and
    (2) Critical center of gravity.
    (c) The maximum limit load rating of each wheel must equal or exceed 
the maximum radial limit load determined under the applicable ground 
load requirements of this part.



Sec. 27.733  Tires.

    (a) Each landing gear wheel must have a tire--
    (1) That is a proper fit on the rim of the wheel; and
    (2) Of the proper rating.
    (b) The maximum static load rating of each tire must equal or exceed 
the static ground reaction obtained at its wheel, assuming--
    (1) The design maximum weight; and
    (2) The most unfavorable center of gravity.
    (c) Each tire installed on a retractable landing gear system must, 
at the maximum size of the tire type expected in service, have a 
clearance to surrounding structure and systems that is adequate to 
prevent contact between the tire and any part of the structure or 
systems.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55469, Dec. 20, 1976]



Sec. 27.735  Brakes.

    For rotorcraft with wheel-type landing gear, a braking device must 
be installed that is--
    (a) Controllable by the pilot;
    (b) Usable during power-off landings; and
    (c) Adequate to--
    (1) Counteract any normal unbalanced torque when starting or 
stopping the rotor; and
    (2) Hold the rotorcraft parked on a 10-degree slope on a dry, smooth 
pavement.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 
49 FR 44434, Nov. 6, 1984]



Sec. 27.737  Skis.

    The maximum limit load rating of each ski must equal or exceed the 
maximum limit load determined under the applicable ground load 
requirements of this part.

                            Floats and Hulls



Sec. 27.751  Main float buoyancy.

    (a) For main floats, the buoyancy necessary to support the maximum 
weight of the rotorcraft in fresh water must be exceeded by--
    (1) 50 percent, for single floats; and
    (2) 60 percent, for multiple floats.
    (b) Each main float must have enough water-tight compartments so 
that, with any single main float compartment flooded, the main floats 
will provide a margin of positive stability great enough to minimize the 
probability of capsizing.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 963, Jan. 26, 1968]



Sec. 27.753  Main float design.

    (a) Bag floats. Each bag float must be designed to withstand--
    (1) The maximum pressure differential that might be developed at the 
maximum altitude for which certification with that float is requested; 
and
    (2) The vertical loads prescribed in Sec. 27.521(a), distributed 
along the length of the bag over three-quarters of its projected area.
    (b) Rigid floats. Each rigid float must be able to withstand the 
vertical, horizontal, and side loads prescribed in Sec. 27.521. These 
loads may be distributed along the length of the float.



Sec. 27.755  Hulls.

    For each rotorcraft, with a hull and auxiliary floats, that is to be 
approved for both taking off from and landing on

[[Page 691]]

water, the hull and auxiliary floats must have enough watertight 
compartments so that, with any single compartment flooded, the buoyancy 
of the hull and auxiliary floats (and wheel tires if used) provides a 
margin of positive stability great enough to minimize the probability of 
capsizing.

                   Personnel and Cargo Accommodations



Sec. 27.771  Pilot compartment.

    For each pilot compartment--
    (a) The compartment and its equipment must allow each pilot to 
perform his duties without unreasonable concentration or fatigue;
    (b) If there is provision for a second pilot, the rotorcraft must be 
controllable with equal safety from either pilot seat; and
    (c) The vibration and noise characteristics of cockpit appurtenances 
may not interfere with safe operation.



Sec. 27.773  Pilot compartment view.

    (a) Each pilot compartment must be free from glare and reflections 
that could interfere with the pilot's view, and designed so that--
    (1) Each pilot's view is sufficiently extensive, clear, and 
undistorted for safe operation; and
    (2) Each pilot is protected from the elements so that moderate rain 
conditions do not unduly impair his view of the flight path in normal 
flight and while landing.
    (b) If certification for night operation is requested, compliance 
with paragraph (a) of this section must be shown in night flight tests.



Sec. 27.775  Windshields and windows.

    Windshields and windows must be made of material that will not break 
into dangerous fragments.

[Amdt. 27-27, 55 FR 38966, Sept. 21, 1990]



Sec. 27.777  Cockpit controls.

    Cockpit controls must be--
    (a) Located to provide convenient operation and to prevent confusion 
and inadvertent operation; and
    (b) Located and arranged with respect to the pilots' seats so that 
there is full and unrestricted movement of each control without 
interference from the cockpit structure or the pilot's clothing when 
pilots from 5[foot]2 to 6[foot]0 in height are 
seated.



Sec. 27.779  Motion and effect of cockpit controls.

    Cockpit controls must be designed so that they operate in accordance 
with the following movements and actuation:
    (a) Flight controls, including the collective pitch control, must 
operate with a sense of motion which corresponds to the effect on the 
rotorcraft.
    (b) Twist-grip engine power controls must be designed so that, for 
lefthand operation, the motion of the pilot's hand is clockwise to 
increase power when the hand is viewed from the edge containing the 
index finger. Other engine power controls, excluding the collective 
control, must operate with a forward motion to increase power.
    (c) Normal landing gear controls must operate downward to extend the 
landing gear.

[Amdt. 27-21, 49 FR 44434, Nov. 6, 1984]



Sec. 27.783  Doors.

    (a) Each closed cabin must have at least one adequate and easily 
accessible external door.
    (b) Each external door must be located where persons using it will 
not be endangered by the rotors, propellers, engine intakes, and 
exhausts when appropriate operating procedures are used. If opening 
procedures are required, they must be marked inside, on or adjacent to 
the door opening device.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-26, 
55 FR 8001, Mar. 6, 1990]



Sec. 27.785  Seats, berths, litters, safety belts, and harnesses.

    (a) Each seat, safety belt, harness, and adjacent part of the 
rotorcraft at each station designated for occupancy during takeoff and 
landing must be free of potentially injurious objects, sharp edges, 
protuberances, and hard surfaces and must be designed so that a person 
making proper use of these facilities will not suffer serious injury in 
an emergency landing as a result of the static inertial load factors 
specified in

[[Page 692]]

Sec. 27.561(b) and dynamic conditions specified in Sec. 27.562.
    (b) Each occupant must be protected from serious head injury by a 
safety belt plus a shoulder harness that will prevent the head from 
contacting any injurious object except as provided for in Sec. 
27.562(c)(5). A shoulder harness (upper torso restraint), in combination 
with the safety belt, constitutes a torso restraint system as described 
in TSO-C114.
    (c) Each occupant's seat must have a combined safety belt and 
shoulder harness with a single-point release. Each pilot's combined 
safety belt and shoulder harness must allow each pilot when seated with 
safety belt and shoulder harness fastened to perform all functions 
necessary for flight operations. There must be a means to secure belts 
and harnesses, when not in use, to prevent interference with the 
operation of the rotorcraft and with rapid egress in an emergency.
    (d) If seat backs do not have a firm handhold, there must be hand 
grips or rails along each aisle to enable the occupants to steady 
themselves while using the aisle in moderately rough air.
    (e) Each projecting object that could injure persons seated or 
moving about in the rotorcraft in normal flight must be padded.
    (f) Each seat and its supporting structure must be designed for an 
occupant weight of at least 170 pounds considering the maximum load 
factors, inertial forces, and reactions between occupant, seat, and 
safety belt or harness corresponding with the applicable flight and 
ground load conditions, including the emergency landing conditions of 
Sec. 27.561(b). In addition--
    (1) Each pilot seat must be designed for the reactions resulting 
from the application of the pilot forces prescribed in Sec. 27.397; and
    (2) The inertial forces prescribed in Sec. 27.561(b) must be 
multiplied by a factor of 1.33 in determining the strength of the 
attachment of--
    (i) Each seat to the structure; and
    (ii) Each safety belt or harness to the seat or structure.
    (g) When the safety belt and shoulder harness are combined, the 
rated strength of the safety belt and shoulder harness may not be less 
than that corresponding to the inertial forces specified in Sec. 
27.561(b), considering the occupant weight of at least 170 pounds, 
considering the dimensional characteristics of the restraint system 
installation, and using a distribution of at least a 60-percent load to 
the safety belt and at least a 40-percent load to the shoulder harness. 
If the safety belt is capable of being used without the shoulder 
harness, the inertial forces specified must be met by the safety belt 
alone.
    (h) When a headrest is used, the headrest and its supporting 
structure must be designed to resist the inertia forces specified in 
Sec. 27.561, with a 1.33 fitting factor and a head weight of at least 
13 pounds.
    (i) Each seating device system includes the device such as the seat, 
the cushions, the occupant restraint system, and attachment devices.
    (j) Each seating device system may use design features such as 
crushing or separation of certain parts of the seats to reduce occupant 
loads for the emergency landing dynamic conditions of Sec. 27.562; 
otherwise, the system must remain intact and must not interfere with 
rapid evacuation of the rotorcraft.
    (k) For the purposes of this section, a litter is defined as a 
device designed to carry a nonambulatory person, primarily in a 
recumbent position, into and on the rotorcraft. Each berth or litter 
must be designed to withstand the load reaction of an occupant weight of 
at least 170 pounds when the occupant is subjected to the forward 
inertial factors specified in Sec. 27.561(b). A berth or litter 
installed within 15[deg] or less of the longitudinal axis of the 
rotorcraft must be provided with a padded end-board, cloth diaphram, or 
equivalent means that can withstand the forward load reaction. A berth 
or litter oriented greater than 15[deg] with the longitudinal axis of 
the rotorcraft must be equipped with appropriate restraints, such as 
straps or safety belts, to withstand the forward load reaction. In 
addition--
    (1) The berth or litter must have a restraint system and must not 
have corners or other protuberances likely to

[[Page 693]]

cause serious injury to a person occupying it during emergency landing 
conditions; and
    (2) The berth or litter attachment and the occupant restraint system 
attachments to the structure must be designed to withstand the critical 
loads resulting from flight and ground load conditions and from the 
conditions prescribed in Sec. 27.561(b). The fitting factor required by 
Sec. 27.625(d) shall be applied.

[Amdt. 27-21, 49 FR 44434, Nov. 6, 1984, as amended by Amdt. 27-25, 54 
FR 47319, Nov. 13, 1989; Amdt. 27-35, 63 FR 43285, Aug. 12, 1998]



Sec. 27.787  Cargo and baggage compartments.

    (a) Each cargo and baggage compartment must be designed for its 
placarded maximum weight of contents and for the critical load 
distributions at the appropriate maximum load factors corresponding to 
the specified flight and ground load conditions, except the emergency 
landing conditions of Sec. 27.561.
    (b) There must be means to prevent the contents of any compartment 
from becoming a hazard by shifting under the loads specified in 
paragraph (a) of this section.
    (c) Under the emergency landing conditions of Sec. 27.561, cargo 
and baggage compartments must--
    (1) Be positioned so that if the contents break loose they are 
unlikely to cause injury to the occupants or restrict any of the escape 
facilities provided for use after an emergency landing; or
    (2) Have sufficient strength to withstand the conditions specified 
in Sec. 27.561 including the means of restraint, and their attachments, 
required by paragraph (b) of this section. Sufficient strength must be 
provided for the maximum authorized weight of cargo and baggage at the 
critical loading distribution.
    (d) If cargo compartment lamps are installed, each lamp must be 
installed so as to prevent contact between lamp bulb and cargo.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55469, Dec. 20, 1976; Amdt. 27-27, 55 FR 38966, Sept. 21, 1990]



Sec. 27.801  Ditching.

    (a) If certification with ditching provisions is requested, the 
rotorcraft must meet the requirements of this section and Sec. Sec. 
27.807(d), 27.1411 and 27.1415.
    (b) Each practicable design measure, compatible with the general 
characteristics of the rotorcraft, must be taken to minimize the 
probability that in an emergency landing on water, the behavior of the 
rotorcraft would cause immediate injury to the occupants or would make 
it impossible for them to escape.
    (c) The probable behavior of the rotorcraft in a water landing must 
be investigated by model tests or by comparison with rotorcraft of 
similar configuration for which the ditching characteristics are known. 
Scoops, flaps, projections, and any other factor likely to affect the 
hydrodynamic characteristics of the rotorcraft must be considered.
    (d) It must be shown that, under reasonably probable water 
conditions, the flotation time and trim of the rotorcraft will allow the 
occupants to leave the rotorcraft and enter the life rafts required by 
Sec. 27.1415. If compliance with this provision is shown by buoyancy 
and trim computations, appropriate allowances must be made for probable 
structural damage and leakage. If the rotorcraft has fuel tanks (with 
fuel jettisoning provisions) that can reasonably be expected to 
withstand a ditching without leakage, the jettisonable volume of fuel 
may be considered as buoyancy volume.
    (e) Unless the effects of the collapse of external doors and windows 
are accounted for in the investigation of the probable behavior of the 
rotorcraft in a water landing (as prescribed in paragraphs (c) and (d) 
of this section), the external doors and windows must be designed to 
withstand the probable maximum local pressures.

[Amdt. 27-11, 41 FR 55469, Dec. 20, 1976]



Sec. 27.805  Flight crew emergency exits.

    (a) For rotorcraft with passenger emergency exits that are not 
convenient to the flight crew, there must be flight crew emergency 
exits, on both

[[Page 694]]

sides of the rotorcraft or as a top hatch in the flight crew area.
    (b) Each flight crew emergency exit must be of sufficient size and 
must be located so as to allow rapid evacuation of the flight crew. This 
must be shown by test.
    (c) Each flight crew emergency exit must not be obstructed by water 
or flotation devices after an emergency landing on water. This must be 
shown by test, demonstration, or analysis.

[Doc. No. 29247, 64 FR 45094, Aug. 18, 1999]



Sec. 27.807  Emergency exits.

    (a) Number and location. (1) There must be at least one emergency 
exit on each side of the cabin readily accessible to each passenger. One 
of these exits must be usable in any probable attitude that may result 
from a crash;
    (2) Doors intended for normal use may also serve as emergency exits, 
provided that they meet the requirements of this section; and
    (3) If emergency flotation devices are installed, there must be an 
emergency exit accessible to each passenger on each side of the cabin 
that is shown by test, demonstration, or analysis to;
    (i) Be above the waterline; and
    (ii) Open without interference from flotation devices, whether 
stowed or deployed.
    (b) Type and operation. Each emergency exit prescribed by paragraph 
(a) of this section must--
    (1) Consist of a movable window or panel, or additional external 
door, providing an unobstructed opening that will admit a 19-by 26-inch 
ellipse;
    (2) Have simple and obvious methods of opening, from the inside and 
from the outside, which do not require exceptional effort;
    (3) Be arranged and marked so as to be readily located and opened 
even in darkness; and
    (4) Be reasonably protected from jamming by fuselage deformation.
    (c) Tests. The proper functioning of each emergency exit must be 
shown by test.
    (d) Ditching emergency exits for passengers. If certification with 
ditching provisions is requested, the markings required by paragraph 
(b)(3) of this section must be designed to remain visible if the 
rotorcraft is capsized and the cabin is submerged.

[Doc. No. 29247, 64 FR 45094, Aug. 18, 1999]



Sec. 27.831  Ventilation.

    (a) The ventilating system for the pilot and passenger compartments 
must be designed to prevent the presence of excessive quantities of fuel 
fumes and carbon monoxide.
    (b) The concentration of carbon monoxide may not exceed one part in 
20,000 parts of air during forward flight or hovering in still air. If 
the concentration exceeds this value under other conditions, there must 
be suitable operating restrictions.



Sec. 27.833  Heaters.

    Each combustion heater must be approved.

[Amdt. 27-23, 53 FR 34210, Sept. 2, 1988]

                             Fire Protection



Sec. 27.853  Compartment interiors.

    For each compartment to be used by the crew or passengers--
    (a) The materials must be at least flame-resistant;
    (b) [Reserved]
    (c) If smoking is to be prohibited, there must be a placard so 
stating, and if smoking is to be allowed--
    (1) There must be an adequate number of self-contained, removable 
ashtrays; and
    (2) Where the crew compartment is separated from the passenger 
compartment, there must be at least one illuminated sign (using either 
letters or symbols) notifying all passengers when smoking is prohibited. 
Signs which notify when smoking is prohibited must--
    (i) When illuminated, be legible to each passenger seated in the 
passenger cabin under all probable lighting conditions; and
    (ii) Be so constructed that the crew can turn the illumination on 
and off.

[Amdt. 27-17, 45 FR 7755, Feb. 4, 1980, as amended by Amdt. 27-37, 64 FR 
45095, Aug. 18, 1999]

[[Page 695]]



Sec. 27.855  Cargo and baggage compartments.

    (a) Each cargo and baggage compartment must be constructed of, or 
lined with, materials that are at least--
    (1) Flame resistant, in the case of compartments that are readily 
accessible to a crewmember in flight; and
    (2) Fire resistant, in the case of other compartments.
    (b) No compartment may contain any controls, wiring, lines, 
equipment, or accessories whose damage or failure would affect safe 
operation, unless those items are protected so that--
    (1) They cannot be damaged by the movement of cargo in the 
compartment; and
    (2) Their breakage or failure will not create a fire hazard.



Sec. 27.859  Heating systems.

    (a) General. For each heating system that involves the passage of 
cabin air over, or close to, the exhaust manifold, there must be means 
to prevent carbon monoxide from entering any cabin or pilot compartment.
    (b) Heat exchangers. Each heat exchanger must be--
    (1) Of suitable materials;
    (2) Adequately cooled under all conditions; and
    (3) Easily disassembled for inspection.
    (c) Combustion heater fire protection. Except for heaters which 
incorporate designs to prevent hazards in the event of fuel leakage in 
the heater fuel system, fire within the ventilating air passage, or any 
other heater malfunction, each heater zone must incorporate the fire 
protection features of the applicable requirements of Sec. Sec. 
27.1183, 27.1185, 27.1189, 27.1191, and be provided with--
    (1) Approved, quick-acting fire detectors in numbers and locations 
ensuring prompt detection of fire in the heater region.
    (2) Fire extinguisher systems that provide at least one adequate 
discharge to all areas of the heater region.
    (3) Complete drainage of each part of each zone to minimize the 
hazards resulting from failure or malfunction of any component 
containing flammable fluids. The drainage means must be--
    (i) Effective under conditions expected to prevail when drainage is 
needed; and
    (ii) Arranged so that no discharged fluid will cause an additional 
fire hazard.
    (4) Ventilation, arranged so that no discharged vapors will cause an 
additional fire hazard.
    (d) Ventilating air ducts. Each ventilating air duct passing through 
any heater region must be fireproof.
    (1) Unless isolation is provided by fireproof valves or by equally 
effective means, the ventilating air duct downstream of each heater must 
be fireproof for a distance great enough to ensure that any fire 
originating in the heater can be contained in the duct.
    (2) Each part of any ventilating duct passing through any region 
having a flammable fluid system must be so constructed or isolated from 
that system that the malfunctioning of any component of that system 
cannot introduce flammable fluids or vapors into the ventilating 
airstream.
    (e) Combustion air ducts. Each combustion air duct must be fireproof 
for a distance great enough to prevent damage from backfiring or reverse 
flame propagation.
    (1) No combustion air duct may connect with the ventilating 
airstream unless flames from backfires or reverse burning cannot enter 
the ventilating airstream under any operating condition, including 
reverse flow or malfunction of the heater or its associated components.
    (2) No combustion air duct may restrict the prompt relief of any 
backfire that, if so restricted, could cause heater failure.
    (f) Heater control: General. There must be means to prevent the 
hazardous accumulation of water or ice on or in any heater control 
component, control system tubing, or safety control.
    (g) Heater safety controls. For each combustion heater, safety 
control means must be provided as follows:
    (1) Means independent of the components provided for the normal 
continuous control of air temperature, airflow, and fuel flow must be 
provided for each heater to automatically shut off the ignition and fuel 
supply of that heater at a point remote from that

[[Page 696]]

heater when any of the following occurs:
    (i) The heat exchanger temperature exceeds safe limits.
    (ii) The ventilating air temperature exceeds safe limits.
    (iii) The combustion airflow becomes inadequate for safe operation.
    (iv) The ventilating airflow becomes inadequate for safe operation.
    (2) The means of complying with paragraph (g)(1) of this section for 
any individual heater must--
    (i) Be independent of components serving any other heater, the heat 
output of which is essential for safe operation; and
    (ii) Keep the heater off until restarted by the crew.
    (3) There must be means to warn the crew when any heater, the heat 
output of which is essential for safe operation, has been shut off by 
the automatic means prescribed in paragraph (g)(1) of this section.
    (h) Air intakes. Each combustion and ventilating air intake must be 
located so that no flammable fluids or vapors can enter the heater 
system--
    (1) During normal operation; or
    (2) As a result of the malfunction of any other component.
    (i) Heater exhaust. Each heater exhaust system must meet the 
requirements of Sec. Sec. 27.1121 and 27.1123.
    (1) Each exhaust shroud must be sealed so that no flammable fluids 
or hazardous quantities of vapors can reach the exhaust system through 
joints.
    (2) No exhaust system may restrict the prompt relief of any backfire 
that, if so restricted, could cause heater failure.
    (j) Heater fuel systems. Each heater fuel system must meet the 
powerplant fuel system requirements affecting safe heater operation. 
Each heater fuel system component in the ventilating airstream must be 
protected by shrouds so that no leakage from those components can enter 
the ventilating airstream.
    (k) Drains. There must be means for safe drainage of any fuel that 
might accumulate in the combustion chamber or the heat exchanger.
    (1) Each part of any drain that operates at high temperatures must 
be protected in the same manner as heater exhausts.
    (2) Each drain must be protected against hazardous ice accumulation 
under any operating condition.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 
53 FR 34211, Sept. 2, 1988]



Sec. 27.861  Fire protection of structure, controls, and other parts.

    Each part of the structure, controls, rotor mechanism, and other 
parts essential to a controlled landing that would be affected by 
powerplant fires must be fireproof or protected so they can perform 
their essential functions for at least 5 minutes under any foreseeable 
powerplant fire conditions.

[Amdt. 27-26, 55 FR 8001, Mar. 6, 1990]



Sec. 27.863  Flammable fluid fire protection.

    (a) In each area where flammable fluids or vapors might escape by 
leakage of a fluid system, there must be means to minimize the 
probability of ignition of the fluids and vapors, and the resultant 
hazards if ignition does occur.
    (b) Compliance with paragraph (a) of this section must be shown by 
analysis or tests, and the following factors must be considered:
    (1) Possible sources and paths of fluid leakage, and means of 
detecting leakage.
    (2) Flammability characteristics of fluids, including effects of any 
combustible or absorbing materials.
    (3) Possible ignition sources, including electrical faults, 
overheating of equipment, and malfunctioning of protective devices.
    (4) Means available for controlling or extinguishing a fire, such as 
stopping flow of fluids, shutting down equipment, fireproof containment, 
or use of extinguishing agents.
    (5) Ability of rotorcraft components that are critical to safety of 
flight to withstand fire and heat.
    (c) If action by the flight crew is required to prevent or 
counteract a fluid fire (e.g. equipment shutdown or actuation of a fire 
extinguisher) quick acting means must be provided to alert the crew.

[[Page 697]]

    (d) Each area where flammable fluids or vapors might escape by 
leakage of a fluid system must be identified and defined.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Amdt. 27-16, 43 FR 50599, Oct. 30, 1978]

                             External Loads



Sec. 27.865  External loads.

    (a) It must be shown by analysis, test, or both, that the rotorcraft 
external load attaching means for rotorcraft-load combinations to be 
used for nonhuman external cargo applications can withstand a limit 
static load equal to 2.5, or some lower load factor approved under 
Sec. Sec. 27.337 through 27.341, multiplied by the maximum external 
load for which authorization is requested. It must be shown by analysis, 
test, or both that the rotorcraft external load attaching means and 
corresponding personnel carrying device system for rotorcraft-load 
combinations to be used for human external cargo applications can 
withstand a limit static load equal to 3.5 or some lower load factor, 
not less than 2.5, approved under Sec. Sec. 27.337 through 27.341, 
multiplied by the maximum external load for which authorization is 
requested. The load for any rotorcraft-load combination class, for any 
external cargo type, must be applied in the vertical direction. For 
jettisonable external loads of any applicable external cargo type, the 
load must also be applied in any direction making the maximum angle with 
the vertical that can be achieved in service but not less than 30[deg]. 
However, the 30[deg] angle may be reduced to a lesser angle if--
    (1) An operating limitation is established limiting external load 
operations to such angles for which compliance with this paragraph has 
been shown; or
    (2) It is shown that the lesser angle can not be exceeded in 
service.
    (b) The external load attaching means, for jettisonable rotorcraft-
load combinations, must include a quick-release system to enable the 
pilot to release the external load quickly during flight. The quick-
release system must consist of a primary quick release subsystem and a 
backup quick release subsystem that are isolated from one another. The 
quick-release system, and the means by which it is controlled, must 
comply with the following:
    (1) A control for the primary quick release subsystem must be 
installed either on one of the pilot's primary controls or in an 
equivalently accessible location and must be designed and located so 
that it may be operated by either the pilot or a crewmember without 
hazardously limiting the ability to control the rotorcraft during an 
emergency situation.
    (2) A control for the backup quick release subsystem, readily 
accessible to either the pilot or another crewmember, must be provided.
    (3) Both the primary and backup quick release subsystems must--
    (i) Be reliable, durable, and function properly with all external 
loads up to and including the maximum external limit load for which 
authorization is requested.
    (ii) Be protected against electromagnetic interference (EMI) from 
external and internal sources and against lightning to prevent 
inadvertent load release.
    (A) The minimum level of protection required for jettisonable 
rotorcraft-load combinations used for nonhuman external cargo is a radio 
frequency field strength of 20 volts per meter.
    (B) The minimum level of protection required for jettisonable 
rotorcraft-load combinations used for human external cargo is a radio 
frequency field strength of 200 volts per meter.
    (iii) Be protected against any failure that could be induced by a 
failure mode of any other electrical or mechanical rotorcraft system.
    (c) For rotorcraft-load combinations to be used for human external 
cargo applications, the rotorcraft must--
    (1) For jettisonable external loads, have a quick-release system 
that meets the requirements of paragraph (b) of this section and that--
    (i) Provides a dual actuation device for the primary quick release 
subsystem, and
    (ii) Provides a separate dual actuation device for the backup quick 
release subsystem;

[[Page 698]]

    (2) Have a reliable, approved personnel carrying device system that 
has the structural capability and personnel safety features essential 
for external occupant safety;
    (3) Have placards and markings at all appropriate locations that 
clearly state the essential system operating instructions and, for the 
personnel carrying device system, the ingress and egress instructions;
    (4) Have equipment to allow direct intercommunication among required 
crewmembers and external occupants; and
    (5) Have the appropriate limitations and procedures incorporated in 
the flight manual for conducting human external cargo operations.
    (d) The critically configured jettisonable external loads must be 
shown by a combination of analysis, ground tests, and flight tests to be 
both transportable and releasable throughout the approved operational 
envelope without hazard to the rotorcraft during normal flight 
conditions. In addition, these external loads must be shown to be 
releasable without hazard to the rotorcraft during emergency flight 
conditions.
    (e) A placard or marking must be installed next to the external-load 
attaching means clearly stating any operational limitations and the 
maximum authorized external load as demonstrated under Sec. 27.25 and 
this section.
    (f) The fatigue evaluation of Sec. 27.571 of this part does not 
apply to rotorcraft-load combinations to be used for nonhuman external 
cargo except for the failure of critical structural elements that would 
result in a hazard to the rotorcraft. For rotorcraft-load combinations 
to be used for human external cargo, the fatigue evaluation of Sec. 
27.571 of this part applies to the entire quick release and personnel 
carrying device structural systems and their attachments.

[Amdt. 27-11, 41 FR 55469, Dec. 20, 1976; as amended by Amdt. 27-26, 55 
FR 8001, Mar. 6, 1990; Amdt. 27-36, 64 FR 43019, Aug. 6, 1999]

                              Miscellaneous



Sec. 27.871  Leveling marks.

    There must be reference marks for leveling the rotorcraft on the 
ground.



Sec. 27.873  Ballast provisions.

    Ballast provisions must be designed and constructed to prevent 
inadvertent shifting of ballast in flight.



                          Subpart E_Powerplant

                                 General



Sec. 27.901  Installation.

    (a) For the purpose of this part, the powerplant installation 
includes each part of the rotorcraft (other than the main and auxiliary 
rotor structures) that--
    (1) Is necessary for propulsion;
    (2) Affects the control of the major propulsive units; or
    (3) Affects the safety of the major propulsive units between normal 
inspections or overhauls.
    (b) For each powerplant installation--
    (1) Each component of the installation must be constructed, 
arranged, and installed to ensure its continued safe operation between 
normal inspections or overhauls for the range of temperature and 
altitude for which approval is requested;
    (2) Accessibility must be provided to allow any inspection and 
maintenance necessary for continued airworthiness;
    (3) Electrical interconnections must be provided to prevent 
differences of potential between major components of the installation 
and the rest of the rotorcraft;
    (4) Axial and radial expansion of turbine engines may not affect the 
safety of the installation; and
    (5) Design precautions must be taken to minimize the possibility of 
incorrect assembly of components and equipment essential to safe 
operation of the rotorcraft, except where operation with the incorrect 
assembly can be shown to be extremely improbable.
    (c) The installation must comply with--
    (1) The installation instructions provided under Sec. 33.5 of this 
chapter; and

[[Page 699]]

    (2) The applicable provisions of this subpart.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 963, Jan. 26, 1968; Amdt. 27-12, 42 FR 15044, Mar. 17, 1977; Amdt. 
27-23, 53 FR 34211, Sept. 2, 1988]



Sec. 27.903  Engines.

    (a) Engine type certification. Each engine must have an approved 
type certificate. Reciprocating engines for use in helicopters must be 
qualified in accordance with Sec. 33.49(d) of this chapter or be 
otherwise approved for the intended usage.
    (b) Engine or drive system cooling fan blade protection. (1) If an 
engine or rotor drive system cooling fan is installed, there must be a 
means to protect the rotorcraft and allow a safe landing if a fan blade 
fails. This must be shown by showing that--
    (i) The fan blades are contained in case of failure;
    (ii) Each fan is located so that a failure will not jeopardize 
safety; or
    (iii) Each fan blade can withstand an ultimate load of 1.5 times the 
centrifugal force resulting from operation limited by the following:
    (A) For fans driven directly by the engine--
    (1) The terminal engine r.p.m. under uncontrolled conditions; or
    (2) An overspeed limiting device.
    (B) For fans driven by the rotor drive system, the maximum rotor 
drive system rotational speed to be expected in service, including 
transients.
    (2) Unless a fatigue evaluation under Sec. 27.571 is conducted, it 
must be shown that cooling fan blades are not operating at resonant 
conditions within the operating limits of the rotorcraft.
    (c) Turbine engine installation. For turbine engine installations, 
the powerplant systems associated with engine control devices, systems, 
and instrumentation must be designed to give reasonable assurance that 
those engine operating limitations that adversely affect turbine rotor 
structural integrity will not be exceeded in service.
    (d) Restart capability: A means to restart any engine in flight must 
be provided.
    (1) Except for the in-flight shutdown of all engines, engine restart 
capability must be demonstrated throughout a flight envelope for the 
rotorcraft.
    (2) Following the in-flight shutdown of all engines, in-flight 
engine restart capability must be provided.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55469, Dec. 20, 1976; Amdt. 27-23, 53 FR 34211, Sept. 2, 1988; 
Amdt. 27-44, 73 FR 11000, Feb. 29, 2008]



Sec. 27.907  Engine vibration.

    (a) Each engine must be installed to prevent the harmful vibration 
of any part of the engine or rotorcraft.
    (b) The addition of the rotor and the rotor drive system to the 
engine may not subject the principal rotating parts of the engine to 
excessive vibration stresses. This must be shown by a vibration 
investigation.
    (c) No part of the rotor drive system may be subjected to excessive 
vibration stresses.

                           Rotor Drive System



Sec. 27.917  Design.

    (a) Each rotor drive system must incorporate a unit for each engine 
to automatically disengage that engine from the main and auxiliary 
rotors if that engine fails.
    (b) Each rotor drive system must be arranged so that each rotor 
necessary for control in autorotation will continue to be driven by the 
main rotors after disengagement of the engine from the main and 
auxiliary rotors.
    (c) If a torque limiting device is used in the rotor drive system, 
it must be located so as to allow continued control of the rotorcraft 
when the device is operating.
    (d) The rotor drive system includes any part necessary to transmit 
power from the engines to the rotor hubs. This includes gear boxes, 
shafting, universal joints, couplings, rotor brake assemblies, clutches, 
supporting bearings for shafting, any attendant accessory pads or 
drives, and any cooling fans that are a part of, attached to, or mounted 
on the rotor drive system.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55469, Dec. 20, 1976]

[[Page 700]]



Sec. 27.921  Rotor brake.

    If there is a means to control the rotation of the rotor drive 
system independently of the engine, any limitations on the use of that 
means must be specified, and the control for that means must be guarded 
to prevent inadvertent operation.



Sec. 27.923  Rotor drive system and control mechanism tests.

    (a) Each part tested as prescribed in this section must be in a 
serviceable condition at the end of the tests. No intervening 
disassembly which might affect test results may be conducted.
    (b) Each rotor drive system and control mechanism must be tested for 
not less than 100 hours. The test must be conducted on the rotorcraft, 
and the torque must be absorbed by the rotors to be installed, except 
that other ground or flight test facilities with other appropriate 
methods of torque absorption may be used if the conditions of support 
and vibration closely simulate the conditions that would exist during a 
test on the rotorcraft.
    (c) A 60-hour part of the test prescribed in paragraph (b) of this 
section must be run at not less than maximum continuous torque and the 
maximum speed for use with maximum continuous torque. In this test, the 
main rotor controls must be set in the position that will give maximum 
longitudinal cyclic pitch change to simulate forward flight. The 
auxiliary rotor controls must be in the position for normal operation 
under the conditions of the test.
    (d) A 30-hour or, for rotorcraft for which the use of either 30-
minute OEI power or continuous OEI power is requested, a 25-hour part of 
the test prescribed in paragraph (b) of this section must be run at not 
less than 75 percent of maximum continuous torque and the minimum speed 
for use with 75 percent of maximum continuous torque. The main and 
auxiliary rotor controls must be in the position for normal operation 
under the conditions of the test.
    (e) A 10-hour part of the test prescribed in paragraph (b) of this 
section must be run at not less than takeoff torque and the maximum 
speed for use with takeoff torque. The main and auxiliary rotor controls 
must be in the normal position for vertical ascent.
    (1) For multiengine rotorcraft for which the use of 2\1/2\ minute 
OEI power is requested, 12 runs during the 10-hour test must be 
conducted as follows:
    (i) Each run must consist of at least one period of 2\1/2\ minutes 
with takeoff torque and the maximum speed for use with takeoff torque on 
all engines.
    (ii) Each run must consist of at least one period for each engine in 
sequence, during which that engine simulates a power failure and the 
remaining engines are run at 2\1/2\ minute OEI torque and the maximum 
speed for use with 2\1/2\ minute OEI torque for 2\1/2\ minutes.
    (2) For multiengine turbine-powered rotorcraft for which the use of 
30-second and 2-minute OEI power is requested, 10 runs must be conducted 
as follows:
    (i) Immediately following a takeoff run of at least 5 minutes, each 
power source must simulate a failure, in turn, and apply the maximum 
torque and the maximum speed for use with 30-second OEI power to the 
remaining affected drive system power inputs for not less than 30 
seconds, followed by application of the maximum torque and the maximum 
speed for use with 2-minute OEI power for not less than 2 minutes. At 
least one run sequence must be conducted from a simulated ``flight 
idle'' condition. When conducted on a bench test, the test sequence must 
be conducted following stabilization at takeoff power.
    (ii) For the purpose of this paragraph, an affected power input 
includes all parts of the rotor drive system which can be adversely 
affected by the application of higher or asymmetric torque and speed 
prescribed by the test.
    (iii) This test may be conducted on a representative bench test 
facility when engine limitations either preclude repeated use of this 
power or would result in premature engine removal during the test. The 
loads, the vibration frequency, and the methods of application to the 
affected rotor drive system components must be representative of 
rotorcraft conditions. Test components must be those used to show 
compliance with the remainder of this section.

[[Page 701]]

    (f) The parts of the test prescribed in paragraphs (c) and (d) of 
this section must be conducted in intervals of not less than 30 minutes 
and may be accomplished either on the ground or in flight. The part of 
the test prescribed in paragraph (e) of this section must be conducted 
in intervals of not less than five minutes.
    (g) At intervals of not more than five hours during the tests 
prescribed in paragraphs (c), (d), and (e) of this section, the engine 
must be stopped rapidly enough to allow the engine and rotor drive to be 
automatically disengaged from the rotors.
    (h) Under the operating conditions specified in paragraph (c) of 
this section, 500 complete cycles of lateral control, 500 complete 
cycles of longitudinal control of the main rotors, and 500 complete 
cycles of control of each auxiliary rotor must be accomplished. A 
``complete cycle'' involves movement of the controls from the neutral 
position, through both extreme positions, and back to the neutral 
position, except that control movements need not produce loads or 
flapping motions exceeding the maximum loads or motions encountered in 
flight. The cycling may be accomplished during the testing prescribed in 
paragraph (c) of this section.
    (i) At least 200 start-up clutch engagements must be accomplished--
    (1) So that the shaft on the driven side of the clutch is 
accelerated; and
    (2) Using a speed and method selected by the applicant.
    (j) For multiengine rotorcraft for which the use of 30-minute OEI 
power is requested, five runs must be made at 30-minute OEI torque and 
the maximum speed for use with 30-minute OEI torque, in which each 
engine, in sequence, is made inoperative and the remaining engine(s) is 
run for a 30-minute period.
    (k) For multiengine rotorcraft for which the use of continuous OEI 
power is requested, five runs must be made at continuous OEI torque and 
the maximum speed for use with continuous OEI torque, in which each 
engine, in sequence, is made inoperative and the remaining engine(s) is 
run for a 1-hour period.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 963, Jan. 26, 1968; Amdt. 27-12, 42 FR 15044, Mar. 17, 1977; Amdt. 
27-23, 53 FR 34212, Sept. 2, 1988; Amdt. 27-29, 59 FR 47767, Sept. 16, 
1994]



Sec. 27.927  Additional tests.

    (a) Any additional dynamic, endurance, and operational tests, and 
vibratory investigations necessary to determine that the rotor drive 
mechanism is safe, must be performed.
    (b) If turbine engine torque output to the transmission can exceed 
the highest engine or transmission torque rating limit, and that output 
is not directly controlled by the pilot under normal operating 
conditions (such as where the primary engine power control is 
accomplished through the flight control), the following test must be 
made:
    (1) Under conditions associated with all engines operating, make 200 
applications, for 10 seconds each, or torque that is at least equal to 
the lesser of--
    (i) The maximum torque used in meeting Sec. 27.923 plus 10 percent; 
or
    (ii) The maximum attainable torque output of the engines, assuming 
that torque limiting devices, if any, function properly.
    (2) For multiengine rotorcraft under conditions associated with each 
engine, in turn, becoming inoperative, apply to the remaining 
transmission torque inputs the maximum torque attainable under probable 
operating conditions, assuming that torque limiting devices, if any, 
function properly. Each transmission input must be tested at this 
maximum torque for at least 15 minutes.
    (3) The tests prescribed in this paragraph must be conducted on the 
rotorcraft at the maximum rotational speed intended for the power 
condition of the test and the torque must be absorbed by the rotors to 
be installed, except that other ground or flight test facilities with 
other appropriate methods of torque absorption may be used if the 
conditions of support and vibration closely simulate the conditions that

[[Page 702]]

would exist during a test on the rotorcraft.
    (c) It must be shown by tests that the rotor drive system is capable 
of operating under autorotative conditions for 15 minutes after the loss 
of pressure in the rotor drive primary oil system.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Amdt. 27-2, 33 FR 963, Jan. 26, 1968, as amended by Amdt. 27-12, 42 FR 
15045, Mar. 17, 1977; Amdt. 27-23, 53 FR 34212, Sept. 2, 1988]



Sec. 27.931  Shafting critical speed.

    (a) The critical speeds of any shafting must be determined by 
demonstration except that analytical methods may be used if reliable 
methods of analysis are available for the particular design.
    (b) If any critical speed lies within, or close to, the operating 
ranges for idling, power on, and autorotative conditions, the stresses 
occurring at that speed must be within safe limits. This must be shown 
by tests.
    (c) If analytical methods are used and show that no critical speed 
lies within the permissible operating ranges, the margins between the 
calculated critical speeds and the limits of the allowable operating 
ranges must be adequate to allow for possible variations between the 
computed and actual values.



Sec. 27.935  Shafting joints.

    Each universal joint, slip joint, and other shafting joints whose 
lubrication is necessary for operation must have provision for 
lubrication.



Sec. 27.939  Turbine engine operating characteristics.

    (a) Turbine engine operating characteristics must be investigated in 
flight to determine that no adverse characteristics (such as stall, 
surge, or flameout) are present, to a hazardous degree, during normal 
and emergency operation within the range of operating limitations of the 
rotorcraft and of the engine.
    (b) The turbine engine air inlet system may not, as a result of 
airflow distortion during normal operation, cause vibration harmful to 
the engine.
    (c) For governor-controlled engines, it must be shown that there 
exists no hazardous torsional instability of the drive system associated 
with critical combinations of power, rotational speed, and control 
displacement.

[Amdt. 27-1, 32 FR 6914, May 5, 1967, as amended by Amdt. 27-11, 41 FR 
55469, Dec. 20, 1976]

                               Fuel System



Sec. 27.951  General.

    (a) Each fuel system must be constructed and arranged to ensure a 
flow of fuel at a rate and pressure established for proper engine 
functioning under any likely operating condition, including the 
maneuvers for which certification is requested.
    (b) Each fuel system must be arranged so that--
    (1) No fuel pump can draw fuel from more than one tank at a time; or
    (2) There are means to prevent introducing air into the system.
    (c) Each fuel system for a turbine engine must be capable of 
sustained operation throughout its flow and pressure range with fuel 
initially saturated with water at 80 [deg]F. and having 0.75cc of free 
water per gallon added and cooled to the most critical condition for 
icing likely to be encountered in operation.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-9, 39 
FR 35461, Oct. 1, 1974]



Sec. 27.952  Fuel system crash resistance.

    Unless other means acceptable to the Administrator are employed to 
minimize the hazard of fuel fires to occupants following an otherwise 
survivable impact (crash landing), the fuel systems must incorporate the 
design features of this section. These systems must be shown to be 
capable of sustaining the static and dynamic deceleration loads of this 
section, considered as ultimate loads acting alone, measured at the 
system component's center of gravity, without structural damage to 
system components, fuel tanks, or their attachments that would leak fuel 
to an ignition source.
    (a) Drop test requirements. Each tank, or the most critical tank, 
must be drop-tested as follows:
    (1) The drop height must be at least 50 feet.
    (2) The drop impact surface must be nondeforming.

[[Page 703]]

    (3) The tank must be filled with water to 80 percent of the normal, 
full capacity.
    (4) The tank must be enclosed in a surrounding structure 
representative of the installation unless it can be established that the 
surrounding structure is free of projections or other design features 
likely to contribute to rupture of the tank.
    (5) The tank must drop freely and impact in a horizontal position 
10[deg].
    (6) After the drop test, there must be no leakage.
    (b) Fuel tank load factors. Except for fuel tanks located so that 
tank rupture with fuel release to either significant ignition sources, 
such as engines, heaters, and auxiliary power units, or occupants is 
extremely remote, each fuel tank must be designed and installed to 
retain its contents under the following ultimate inertial load factors, 
acting alone.
    (1) For fuel tanks in the cabin:
    (i) Upward--4g.
    (ii) Forward--16g.
    (iii) Sideward--8g.
    (iv) Downward--20g.
    (2) For fuel tanks located above or behind the crew or passenger 
compartment that, if loosened, could injure an occupant in an emergency 
landing:
    (i) Upward--1.5g.
    (ii) Forward--8g.
    (iii) Sideward--2g.
    (iv) Downward--4g.
    (3) For fuel tanks in other areas:
    (i) Upward--1.5g.
    (ii) Forward--4g.
    (iii) Sideward--2g.
    (iv) Downward--4g.
    (c) Fuel line self-sealing breakaway couplings. Self-sealing 
breakaway couplings must be installed unless hazardous relative motion 
of fuel system components to each other or to local rotorcraft structure 
is demonstrated to be extremely improbable or unless other means are 
provided. The couplings or equivalent devices must be installed at all 
fuel tank-to-fuel line connections, tank-to-tank interconnects, and at 
other points in the fuel system where local structural deformation could 
lead to the release of fuel.
    (1) The design and construction of self-sealing breakaway couplings 
must incorporate the following design features:
    (i) The load necessary to separate a breakaway coupling must be 
between 25 to 50 percent of the minimum ultimate failure load (ultimate 
strength) of the weakest component in the fluid-carrying line. The 
separation load must in no case be less than 300 pounds, regardless of 
the size of the fluid line.
    (ii) A breakaway coupling must separate whenever its ultimate load 
(as defined in paragraph (c)(1)(i) of this section) is applied in the 
failure modes most likely to occur.
    (iii) All breakaway couplings must incorporate design provisions to 
visually ascertain that the coupling is locked together (leak-free) and 
is open during normal installation and service.
    (iv) All breakaway couplings must incorporate design provisions to 
prevent uncoupling or unintended closing due to operational shocks, 
vibrations, or accelerations.
    (v) No breakaway coupling design may allow the release of fuel once 
the coupling has performed its intended function.
    (2) All individual breakaway couplings, coupling fuel feed systems, 
or equivalent means must be designed, tested, installed, and maintained 
so that inadvertent fuel shutoff in flight is improbable in accordance 
with Sec. 27.955(a) and must comply with the fatigue evaluation 
requirements of Sec. 27.571 without leaking.
    (3) Alternate, equivalent means to the use of breakaway couplings 
must not create a survivable impact-induced load on the fuel line to 
which it is installed greater than 25 to 50 percent of the ultimate load 
(strength) of the weakest component in the line and must comply with the 
fatigue requirements of Sec. 27.571 without leaking.
    (d) Frangible or deformable structural attachments. Unless hazardous 
relative motion of fuel tanks and fuel system components to local 
rotorcraft structure is demonstrated to be extremely improbable in an 
otherwise survivable impact, frangible or locally deformable attachments 
of fuel tanks and fuel system components to local rotorcraft structure 
must be used. The attachment of fuel tanks and fuel system

[[Page 704]]

components to local rotorcraft structure, whether frangible or locally 
deformable, must be designed such that its separation or relative local 
deformation will occur without rupture or local tear-out of the fuel 
tank or fuel system components that will cause fuel leakage. The 
ultimate strength of frangible or deformable attachments must be as 
follows:
    (1) The load required to separate a frangible attachment from its 
support structure, or deform a locally deformable attachment relative to 
its support structure, must be between 25 and 50 percent of the minimum 
ultimate load (ultimate strength) of the weakest component in the 
attached system. In no case may the load be less than 300 pounds.
    (2) A frangible or locally deformable attachment must separate or 
locally deform as intended whenever its ultimate load (as defined in 
paragraph (d)(1) of this section) is applied in the modes most likely to 
occur.
    (3) All frangible or locally deformable attachments must comply with 
the fatigue requirements of Sec. 27.571.
    (e) Separation of fuel and ignition sources. To provide maximum 
crash resistance, fuel must be located as far as practicable from all 
occupiable areas and from all potential ignition sources.
    (f) Other basic mechanical design criteria. Fuel tanks, fuel lines, 
electrical wires, and electrical devices must be designed, constructed, 
and installed, as far as practicable, to be crash resistant.
    (g) Rigid or semirigid fuel tanks. Rigid or semirigid fuel tank or 
bladder walls must be impact and tear resistant.

[Doc. No. 26352, 59 FR 50386, Oct. 3, 1994]



Sec. 27.953  Fuel system independence.

    (a) Each fuel system for multiengine rotorcraft must allow fuel to 
be supplied to each engine through a system independent of those parts 
of each system supplying fuel to other engines. However, separate fuel 
tanks need not be provided for each engine.
    (b) If a single fuel tank is used on a multiengine rotorcraft, the 
following must be provided:
    (1) Independent tank outlets for each engine, each incorporating a 
shutoff valve at the tank. This shutoff valve may also serve as the 
firewall shutoff valve required by Sec. 27.995 if the line between the 
valve and the engine compartment does not contain a hazardous amount of 
fuel that can drain into the engine compartment.
    (2) At least two vents arranged to minimize the probability of both 
vents becoming obstructed simultaneously.
    (3) Filler caps designed to minimize the probability of incorrect 
installation or inflight loss.
    (4) A fuel system in which those parts of the system from each tank 
outlet to any engine are independent of each part of each system 
supplying fuel to other engines.



Sec. 27.954  Fuel system lightning protection.

    The fuel system must be designed and arranged to prevent the 
ignition of fuel vapor within the system by--
    (a) Direct lightning strikes to areas having a high probability of 
stroke attachment;
    (b) Swept lightning strokes to areas where swept strokes are highly 
probable; or
    (c) Corona and streamering at fuel vent outlets.

[Amdt. 27-23, 53 FR 34212, Sept. 2, 1988]



Sec. 27.955  Fuel flow.

    (a) General. The fuel system for each engine must be shown to 
provide the engine with at least 100 percent of the fuel required under 
each operating and maneuvering condition to be approved for the 
rotorcraft including, as applicable, the fuel required to operate the 
engine(s) under the test conditions required by Sec. 27.927. Unless 
equivalent methods are used, compliance must be shown by test during 
which the following provisions are met except that combinations of 
conditions which are shown to be improbable need not be considered.
    (1) The fuel pressure, corrected for critical accelerations, must be 
within the limits specified by the engine type certificate data sheet.
    (2) The fuel level in the tank may not exceed that established as 
the unusable fuel supply for that tank under Sec. 27.959, plus the 
minimum additional fuel necessary to conduct the test.

[[Page 705]]

    (3) The fuel head between the tank outlet and the engine inlet must 
be critical with respect to rotorcraft flight attitudes.
    (4) The critical fuel pump (for pump-fed systems) is installed to 
produce (by actual or simulated failure) the critical restriction to 
fuel flow to be expected from pump failure.
    (5) Critical values of engine rotation speed, electrical power, or 
other sources of fuel pump motive power must be applied.
    (6) Critical values of fuel properties which adversely affect fuel 
flow must be applied.
    (7) The fuel filter required by Sec. 27.997 must be blocked to the 
degree necessary to simulate the accumulation of fuel contamination 
required to activate the indicator required by Sec. 27.1305(q).
    (b) Fuel transfer systems. If normal operation of the fuel system 
requires fuel to be transferred to an engine feed tank, the transfer 
must occur automatically via a system which has been shown to maintain 
the fuel level in the engine feed tank within acceptable limits during 
flight or surface operation of the rotorcraft.
    (c) Multiple fuel tanks. If an engine can be supplied with fuel from 
more than one tank, the fuel systems must, in addition to having 
appropriate manual switching capability, be designed to prevent 
interruption of fuel flow to that engine, without attention by the 
flightcrew, when any tank supplying fuel to that engine is depleted of 
usable fuel during normal operation, and any other tank that normally 
supplies fuel to the engine alone contains usable fuel.

[Amdt. 27-23, 53 FR 34212, Sept. 2, 1988]



Sec. 27.959  Unusable fuel supply.

    The unusable fuel supply for each tank must be established as not 
less than the quantity at which the first evidence of malfunction occurs 
under the most adverse fuel feed condition occurring under any intended 
operations and flight maneuvers involving that tank.



Sec. 27.961  Fuel system hot weather operation.

    Each suction lift fuel system and other fuel systems with features 
conducive to vapor formation must be shown by test to operate 
satisfactorily (within certification limits) when using fuel at a 
temperature of 110 [deg]F under critical operating conditions including, 
if applicable, the engine operating conditions defined by Sec. 27.927 
(b)(1) and (b)(2).

[Amdt. 27-23, 53 FR 34212, Sept. 2, 1988]



Sec. 27.963  Fuel tanks: general.

    (a) Each fuel tank must be able to withstand, without failure, the 
vibration, inertia, fluid, and structural loads to which it may be 
subjected in operation.
    (b) Each fuel tank of 10 gallons or greater capacity must have 
internal baffles, or must have external support to resist surging.
    (c) Each fuel tank must be separated from the engine compartment by 
a firewall. At least one-half inch of clear airspace must be provided 
between the tank and the firewall.
    (d) Spaces adjacent to the surfaces of fuel tanks must be ventilated 
so that fumes cannot accumulate in the tank compartment in case of 
leakage. If two or more tanks have interconnected outlets, they must be 
considered as one tank, and the airspaces in those tanks must be 
interconnected to prevent the flow of fuel from one tank to another as a 
result of a difference in pressure between those airspaces.
    (e) The maximum exposed surface temperature of any component in the 
fuel tank must be less, by a safe margin as determined by the 
Administrator, than the lowest expected autoignition temperature of the 
fuel or fuel vapor in the tank. Compliance with this requirement must be 
shown under all operating conditions and under all failure or 
malfunction conditions of all components inside the tank.
    (f) Each fuel tank installed in personnel compartments must be 
isolated by fume-proof and fuel-proof enclosures that are drained and 
vented to the exterior of the rotorcraft. The design and construction of 
the enclosures must provide necessary protection for the

[[Page 706]]

tank, must be crash resistant during a survivable impact in accordance 
with Sec. 27.952, and must be adequate to withstand loads and abrasions 
to be expected in personnel compartments.
    (g) Each flexible fuel tank bladder or liner must be approved or 
shown to be suitable for the particular application and must be puncture 
resistant. Puncture resistance must be shown by meeting the TSO-C80, 
paragraph 16.0, requirements using a minimum puncture force of 370 
pounds.
    (h) Each integral fuel tank must have provisions for inspection and 
repair of its interior.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 
53 FR 34213, Sept. 2, 1988; Amdt. 27-30, 59 FR 50387, Oct. 3, 1994]



Sec. 27.965  Fuel tank tests.

    (a) Each fuel tank must be able to withstand the applicable pressure 
tests in this section without failure or leakage. If practicable, test 
pressures may be applied in a manner simulating the pressure 
distribution in service.
    (b) Each conventional metal tank, nonmetallic tank with walls that 
are not supported by the rotorcraft structure, and integral tank must be 
subjected to a pressure of 3.5 p.s.i. unless the pressure developed 
during maximum limit acceleration or emergency deceleration with a full 
tank exceeds this value, in which case a hydrostatic head, or equivalent 
test, must be applied to duplicate the acceleration loads as far as 
possible. However, the pressure need not exceed 3.5 p.s.i. on surfaces 
not exposed to the acceleration loading.
    (c) Each nonmetallic tank with walls supported by the rotorcraft 
structure must be subjected to the following tests:
    (1) A pressure test of at least 2.0 p.s.i. This test may be 
conducted on the tank alone in conjunction with the test specified in 
paragraph (c)(2) of this section.
    (2) A pressure test, with the tank mounted in the rotorcraft 
structure, equal to the load developed by the reaction of the contents, 
with the tank full, during maximum limit acceleration or emergency 
deceleration. However, the pressure need not exceed 2.0 p.s.i. on 
surfaces not exposed to the acceleration loading.
    (d) Each tank with large unsupported or unstiffened flat areas, or 
with other features whose failure or deformation could cause leakage, 
must be subjected to the following test or its equivalent:
    (1) Each complete tank assembly and its support must be vibration 
tested while mounted to simulate the actual installation.
    (2) The tank assembly must be vibrated for 25 hours while two-thirds 
full of any suitable fluid. The amplitude of vibration may not be less 
than one thirty-second of an inch, unless otherwise substantiated.
    (3) The test frequency of vibration must be as follows:
    (i) If no frequency of vibration resulting from any r.p.m. within 
the normal operating range of engine or rotor system speeds is critical, 
the test frequency of vibration, in number of cycles per minute must, 
unless a frequency based on a more rational calculation is used, be the 
number obtained by averaging the maximum and minimum power-on engine 
speeds (r.p.m.) for reciprocating engine powered rotorcraft or 2,000 
c.p.m. for turbine engine powered rotorcraft.
    (ii) If only one frequency of vibration resulting from any r.p.m. 
within the normal operating range of engine or rotor system speeds is 
critical, that frequency of vibration must be the test frequency.
    (iii) If more than one frequency of vibration resulting from any 
r.p.m. within the normal operating range of engine or rotor system 
speeds is critical, the most critical of these frequencies must be the 
test frequency.
    (4) Under paragraphs (d)(3)(ii) and (iii) of this section, the time 
of test must be adjusted to accomplish the same number of vibration 
cycles as would be accomplished in 25 hours at the frequency specified 
in paragraph (d)(3)(i) of this section.
    (5) During the test, the tank assembly must be rocked at the rate of 
16 to 20 complete cycles per minute through an angle of 15 degrees on 
both sides of the horizontal (30 degrees total), about the most critical 
axis, for 25 hours. If motion about more than one axis is likely to be 
critical, the tank must be

[[Page 707]]

rocked about each critical axis for 12\1/2\ hours.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Amdt. 27-12, 42 FR 15045, Mar. 17, 1977]



Sec. 27.967  Fuel tank installation.

    (a) Each fuel tank must be supported so that tank loads are not 
concentrated on unsupported tank surfaces. In addition--
    (1) There must be pads, if necessary, to prevent chafing between 
each tank and its supports;
    (2) The padding must be nonabsorbent or treated to prevent the 
absorption of fuel;
    (3) If flexible tank liners are used, they must be supported so that 
it is not necessary for them to withstand fluid loads; and
    (4) Each interior surface of tank compartments must be smooth and 
free of projections that could cause wear of the liner unless--
    (i) There are means for protection of the liner at those points; or
    (ii) The construction of the liner itself provides such protection.
    (b) Any spaces adjacent to tank surfaces must be adequately 
ventilated to avoid accumulation of fuel or fumes in those spaces due to 
minor leakage. If the tank is in a sealed compartment, ventilation may 
be limited to drain holes that prevent clogging and excessive pressure 
resulting from altitude changes. If flexible tank liners are installed, 
the venting arrangement for the spaces between the liner and its 
container must maintain the proper relationship to tank vent pressures 
for any expected flight condition.
    (c) The location of each tank must meet the requirements of Sec. 
27.1185 (a) and (c).
    (d) No rotorcraft skin immediately adjacent to a major air outlet 
from the engine compartment may act as the wall of the integral tank.

[Doc. No. 26352, 59 FR 50387, Oct. 3, 1994]



Sec. 27.969  Fuel tank expansion space.

    Each fuel tank or each group of fuel tanks with interconnected vent 
systems must have an expansion space of not less than 2 percent of the 
tank capacity. It must be impossible to fill the fuel tank expansion 
space inadvertently with the rotorcraft in the normal ground attitude.

[Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]



Sec. 27.971  Fuel tank sump.

    (a) Each fuel tank must have a drainable sump with an effective 
capacity in any ground attitude to be expected in service of 0.25 
percent of the tank capacity or \1/16\ gallon, whichever is greater, 
unless--
    (1) The fuel system has a sediment bowl or chamber that is 
accessible for preflight drainage and has a minimum capacity of 1 ounce 
for every 20 gallons of fuel tank capacity; and
    (2) Each fuel tank drain is located so that in any ground attitude 
to be expected in service, water will drain from all parts of the tank 
to the sediment bowl or chamber.
    (b) Each sump, sediment bowl, and sediment chamber drain required by 
this section must comply with the drain provisions of Sec. 27.999(b).

[Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]



Sec. 27.973  Fuel tank filler connection.

    (a) Each fuel tank filler connection must prevent the entrance of 
fuel into any part of the rotorcraft other than the tank itself during 
normal operations and must be crash resistant during a survivable impact 
in accordance with Sec. 27.952(c). In addition--
    (1) Each filler must be marked as prescribed in Sec. 27.1557(c)(1);
    (2) Each recessed filler connection that can retain any appreciable 
quantity of fuel must have a drain that discharges clear of the entire 
rotorcraft; and
    (3) Each filler cap must provide a fuel-tight seal under the fluid 
pressure expected in normal operation and in a survivable impact.
    (b) Each filler cap or filler cap cover must warn when the cap is 
not fully locked or seated on the filler connection.

[Doc. No. 26352, 59 FR 50387, Oct. 3, 1994]



Sec. 27.975  Fuel tank vents.

    (a) Each fuel tank must be vented from the top part of the expansion 
space so that venting is effective under

[[Page 708]]

all normal flight conditions. Each vent must minimize the probability of 
stoppage by dirt or ice.
    (b) The venting system must be designed to minimize spillage of fuel 
through the vents to an ignition source in the event of a rollover 
during landing, ground operation, or a survivable impact.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 
53 FR 34213, Sept. 2, 1988; Amdt. 27-30, 59 FR 50387, Oct. 3, 1994; 
Amdt. 27-35, 63 FR 43285, Aug. 12, 1998]



Sec. 27.977  Fuel tank outlet.

    (a) There must be a fuel stainer for the fuel tank outlet or for the 
booster pump. This strainer must--
    (1) For reciprocating engine powered rotorcraft, have 8 to 16 meshes 
per inch; and
    (2) For turbine engine powered rotorcraft, prevent the passage of 
any object that could restrict fuel flow or damage any fuel system 
component.
    (b) The clear area of each fuel tank outlet strainer must be at 
least five times the area of the outlet line.
    (c) The diameter of each strainer must be at least that of the fuel 
tank outlet.
    (d) Each finger strainer must be accessible for inspection and 
cleaning.

[Amdt. 27-11, 41 FR 55470, Dec. 20, 1976]

                         Fuel System Components



Sec. 27.991  Fuel pumps.

    Compliance with Sec. 27.955 may not be jeopardized by failure of--
    (a) Any one pump except pumps that are approved and installed as 
parts of a type certificated engine; or
    (b) Any component required for pump operation except, for engine 
driven pumps, the engine served by that pump.

[Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]



Sec. 27.993  Fuel system lines and fittings.

    (a) Each fuel line must be installed and supported to prevent 
excessive vibration and to withstand loads due to fuel pressure and 
accelerated flight conditions.
    (b) Each fuel line connected to components of the rotorcraft between 
which relative motion could exist must have provisions for flexibility.
    (c) Flexible hose must be approved.
    (d) Each flexible connection in fuel lines that may be under 
pressure or subjected to axial loading must use flexible hose 
assemblies.
    (e) No flexible hose that might be adversely affected by high 
temperatures may be used where excessive temperatures will exist during 
operation or after engine shutdown.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 964, Jan. 26, 1968]



Sec. 27.995  Fuel valves.

    (a) There must be a positive, quick-acting valve to shut off fuel to 
each engine individually.
    (b) The control for this valve must be within easy reach of 
appropriate crewmembers.
    (c) Where there is more than one source of fuel supply there must be 
means for independent feeding from each source.
    (d) No shutoff valve may be on the engine side of any firewall.



Sec. 27.997  Fuel strainer or filter.

    There must be a fuel strainer or filter between the fuel tank outlet 
and the inlet of the first fuel system component which is susceptible to 
fuel contamination, including but not limited to the fuel metering 
device or an engine positive displacement pump, whichever is nearer the 
fuel tank outlet. This fuel strainer or filter must--
    (a) Be accessible for draining and cleaning and must incorporate a 
screen or element which is easily removable;
    (b) Have a sediment trap and drain except that it need not have a 
drain if the strainer or filter is easily removable for drain purposes;
    (c) Be mounted so that its weight is not supported by the connecting 
lines or by the inlet or outlet connections of the strainer or filter 
itself, unless adequate strength margins under all loading conditions 
are provided in the lines and connections; and
    (d) Provide a means to remove from the fuel any contaminant which 
would jeopardize the flow of fuel through rotorcraft or engine fuel 
system components required for proper rotorcraft

[[Page 709]]

fuel system or engine fuel system operation.

[Amdt. 27-9, 39 FR 35461, Oct. 1, 1974, as amended by Amdt. 27-20, 49 FR 
6849, Feb. 23, 1984; Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]



Sec. 27.999  Fuel system drains.

    (a) There must be at least one accessible drain at the lowest point 
in each fuel system to completely drain the system with the rotorcraft 
in any ground attitude to be expected in service.
    (b) Each drain required by paragraph (a) of this section must--
    (1) Discharge clear of all parts of the rotorcraft;
    (2) Have manual or automatic means to assure positive closure in the 
off position; and
    (3) Have a drain valve--
    (i) That is readily accessible and which can be easily opened and 
closed; and
    (ii) That is either located or protected to prevent fuel spillage in 
the event of a landing with landing gear retracted.

[Doc. No. 574, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 
FR 55470, Dec. 20, 1976; Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]

                               Oil System



Sec. 27.1011  Engines: General.

    (a) Each engine must have an independent oil system that can supply 
it with an appropriate quantity of oil at a temperature not above that 
safe for continuous operation.
    (b) The usable oil capacity of each system may not be less than the 
product of the endurance of the rotorcraft under critical operating 
conditions and the maximum oil consumption of the engine under the same 
conditions, plus a suitable margin to ensure adequate circulation and 
cooling. Instead of a rational analysis of endurance and consumption, a 
usable oil capacity of one gallon for each 40 gallons of usable fuel may 
be used.
    (c) The oil cooling provisions for each engine must be able to 
maintain the oil inlet temperature to that engine at or below the 
maximum established value. This must be shown by flight tests.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 
53 FR 34213, Sept. 2, 1988]



Sec. 27.1013  Oil tanks.

    Each oil tank must be designed and installed so that--
    (a) It can withstand, without failure, each vibration, inertia, 
fluid, and structural load expected in operation;
    (b) [Reserved]
    (c) Where used with a reciprocating engine, it has an expansion 
space of not less than the greater of 10 percent of the tank capacity or 
0.5 gallon, and where used with a turbine engine, it has an expansion 
space of not less than 10 percent of the tank capacity.
    (d) It is impossible to fill the tank expansion space inadvertently 
with the rotorcraft in the normal ground attitude;
    (e) Adequate venting is provided; and
    (f) There are means in the filler opening to prevent oil overflow 
from entering the oil tank compartment.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-9, 39 
FR 35461, Oct. 1, 1974]



Sec. 27.1015  Oil tank tests.

    Each oil tank must be designed and installed so that it can 
withstand, without leakage, an internal pressure of 5 p.s.i., except 
that each pressurized oil tank used with a turbine engine must be 
designed and installed so that it can withstand, without leakage, an 
internal pressure of 5 p.s.i., plus the maximum operating pressure of 
the tank.

[Amdt. 27-9, 39 FR 35462, Oct. 1, 1974]



Sec. 27.1017  Oil lines and fittings.

    (a) Each oil line must be supported to prevent excessive vibration.
    (b) Each oil line connected to components of the rotorcraft between 
which relative motion could exist must have provisions for flexibility.
    (c) Flexible hose must be approved.
    (d) Each oil line must have an inside diameter of not less than the 
inside diameter of the engine inlet or outlet. No line may have splices 
between connections.

[[Page 710]]



Sec. 27.1019  Oil strainer or filter.

    (a) Each turbine engine installation must incorporate an oil 
strainer or filter through which all of the engine oil flows and which 
meets the following requirements:
    (1) Each oil strainer or filter that has a bypass must be 
constructed and installed so that oil will flow at the normal rate 
through the rest of the system with the strainer or filter completely 
blocked.
    (2) The oil strainer or filter must have the capacity (with respect 
to operating limitations established for the engine) to ensure that 
engine oil system functioning is not impaired when the oil is 
contaminated to a degree (with respect to particle size and density) 
that is greater than that established for the engine under Part 33 of 
this chapter.
    (3) The oil strainer or filter, unless it is installed at an oil 
tank outlet, must incorporate a means to indicate contamination before 
it reaches the capacity established in accordance with paragraph (a)(2) 
of this section.
    (4) The bypass of a strainer or filter must be constructed and 
installed so that the release of collected contaminants is minimized by 
appropriate location of the bypass to ensure that collected contaminants 
are not in the bypass flow path.
    (5) An oil strainer or filter that has no bypass, except one that is 
installed at an oil tank outlet, must have a means to connect it to the 
warning system required in Sec. 27.1305(r).
    (b) Each oil strainer or filter in a powerplant installation using 
reciprocating engines must be constructed and installed so that oil will 
flow at the normal rate through the rest of the system with the strainer 
or filter element completely blocked.

[Amdt. 27-9, 39 FR 35462, Oct. 1, 1974, as amended by Amdt. 27-20, 49 FR 
6849, Feb. 23, 1984; Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]



Sec. 27.1021  Oil system drains.

    A drain (or drains) must be provided to allow safe drainage of the 
oil system. Each drain must--
    (a) Be accessible; and
    (b) Have manual or automatic means for positive locking in the 
closed position.

[Amdt. 27-20, 49 FR 6849, Feb. 23, 1984]



Sec. 27.1027  Transmissions and gearboxes: General.

    (a) The lubrication system for components of the rotor drive system 
that require continuous lubrication must be sufficiently independent of 
the lubrication systems of the engine(s) to ensure lubrication during 
autorotation.
    (b) Pressure lubrication systems for transmissions and gearboxes 
must comply with the engine oil system requirements of Sec. Sec. 
27.1013 (except paragraph (c)), 27.1015, 27.1017, 27.1021, and 
27.1337(d).
    (c) Each pressure lubrication system must have an oil strainer or 
filter through which all of the lubricant flows and must--
    (1) Be designed to remove from the lubricant any contaminant which 
may damage transmission and drive system components or impede the flow 
of lubricant to a hazardous degree;
    (2) Be equipped with a means to indicate collection of contaminants 
on the filter or strainer at or before opening of the bypass required by 
paragraph (c)(3) of this section; and
    (3) Be equipped with a bypass constructed and installed so that--
    (i) The lubricant will flow at the normal rate through the rest of 
the system with the strainer or filter completely blocked; and
    (ii) The release of collected contaminants is minimized by 
appropriate location of the bypass to ensure that collected contaminants 
are not in the bypass flowpath.
    (d) For each lubricant tank or sump outlet supplying lubrication to 
rotor drive systems and rotor drive system components, a screen must be 
provided to prevent entrance into the lubrication system of any object 
that might obstruct the flow of lubricant from the outlet to the filter 
required by paragraph (c) of this section. The requirements of paragraph 
(c) do not apply to screens installed at lubricant tank or sump outlets.

[[Page 711]]

    (e) Splash-type lubrication systems for rotor drive system gearboxes 
must comply with Sec. Sec. 27.1021 and 27.1337(d).

[Amdt. 27-23, 53 FR 34213, Sept. 2, 1988, as amended by Amdt. 27-37, 64 
FR 45095, Aug. 18, 1999]

                                 Cooling



Sec. 27.1041  General.

    (a) Each powerplant cooling system must be able to maintain the 
temperatures of powerplant components within the limits established for 
these components under critical surface (ground or water) and flight 
operating conditions for which certification is required and after 
normal shutdown. Powerplant components to be considered include but may 
not be limited to engines, rotor drive system components, auxiliary 
power units, and the cooling or lubricating fluids used with these 
components.
    (b) Compliance with paragraph (a) of this section must be shown in 
tests conducted under the conditions prescribed in that paragraph.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 
53 FR 34213, Sept. 2, 1988]



Sec. 27.1043  Cooling tests.

    (a) General. For the tests prescribed in Sec. 27.1041(b), the 
following apply:
    (1) If the tests are conducted under conditions deviating from the 
maximum ambient atmospheric temperature specified in paragraph (b) of 
this section, the recorded powerplant temperatures must be corrected 
under paragraphs (c) and (d) of this section unless a more rational 
correction method is applicable.
    (2) No corrected temperature determined under paragraph (a)(1) of 
this section may exceed established limits.
    (3) For reciprocating engines, the fuel used during the cooling 
tests must be of the minimum grade approved for the engines, and the 
mixture settings must be those normally used in the flight stages for 
which the cooling tests are conducted.
    (4) The test procedures must be as prescribed in Sec. 27.1045.
    (b) Maximum ambient atmospheric temperature. A maximum ambient 
atmospheric temperature corresponding to sea level conditions of at 
least 100 degrees F. must be established. The assumed temperature lapse 
rate is 3.6 degrees F. per thousand feet of altitude above sea level 
until a temperature of -69.7 degrees F. is reached, above which altitude 
the temperature is considered constant at -69.7 degrees F. However, for 
winterization installations, the applicant may select a maximum ambient 
atmospheric temperature corresponding to sea level conditions of less 
than 100 degrees F.
    (c) Correction factor (except cylinder barrels). Unless a more 
rational correction applies, temperatures of engine fluids and power-
plant components (except cylinder barrels) for which temperature limits 
are established, must be corrected by adding to them the difference 
between the maximum ambient atmospheric temperature and the temperature 
of the ambient air at the time of the first occurrence of the maximum 
component or fluid temperature recorded during the cooling test.
    (d) Correction factor for cylinder barrel temperatures. Cylinder 
barrel temperatures must be corrected by adding to them 0.7 times the 
difference between the maximum ambient atmospheric temperature and the 
temperature of the ambient air at the time of the first occurrence of 
the maximum cylinder barrel temperature recorded during the cooling 
test.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55470, Dec. 20, 1976; Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]



Sec. 27.1045  Cooling test procedures.

    (a) General. For each stage of flight, the cooling tests must be 
conducted with the rotorcraft--
    (1) In the configuration most critical for cooling; and
    (2) Under the conditions most critical for cooling.
    (b) Temperature stabilization. For the purpose of the cooling tests, 
a temperature is ``stabilized'' when its rate of change is less than two 
degrees F. per minute. The following component and

[[Page 712]]

engine fluid temperature stabilization rules apply:
    (1) For each rotorcraft, and for each stage of flight--
    (i) The temperatures must be stabilized under the conditions from 
which entry is made into the stage of flight being investigated; or
    (ii) If the entry condition normally does not allow temperatures to 
stabilize, operation through the full entry condition must be conducted 
before entry into the stage of flight being investigated in order to 
allow the temperatures to attain their natural levels at the time of 
entry.
    (2) For each helicopter during the takeoff stage of flight, the 
climb at takeoff power must be preceded by a period of hover during 
which the temperatures are stabilized.
    (c) Duration of test. For each stage of flight the tests must be 
continued until--
    (1) The temperatures stabilize or 5 minutes after the occurrence of 
the highest temperature recorded, as appropriate to the test condition;
    (2) That stage of flight is completed; or
    (3) An operating limitation is reached.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 
53 FR 34214, Sept. 2, 1988]

                            Induction System



Sec. 27.1091  Air induction.

    (a) The air induction system for each engine must supply the air 
required by that engine under the operating conditions and maneuvers for 
which certification is requested.
    (b) Each cold air induction system opening must be outside the 
cowling if backfire flames can emerge.
    (c) If fuel can accumulate in any air induction system, that system 
must have drains that discharge fuel--
    (1) Clear of the rotorcraft; and
    (2) Out of the path of exhaust flames.
    (d) For turbine engine powered rotorcraft--
    (1) There must be means to prevent hazardous quantities of fuel 
leakage or overflow from drains, vents, or other components of flammable 
fluid systems from entering the engine intake system; and
    (2) The air inlet ducts must be located or protected so as to 
minimize the ingestion of foreign matter during takeoff, landing, and 
taxiing.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 964, Jan. 26, 1968; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988]



Sec. 27.1093  Induction system icing protection.

    (a) Reciprocating engines. Each reciprocating engine air induction 
system must have means to prevent and eliminate icing. Unless this is 
done by other means, it must be shown that, in air free of visible 
moisture at a temperature of 30 degrees F., and with the engines at 75 
percent of maximum continuous power--
    (1) Each rotorcraft with sea level engines using conventional 
venturi carburetors has a preheater that can provide a heat rise of 90 
degrees F.;
    (2) Each rotorcraft with sea level engines using carburetors tending 
to prevent icing has a sheltered alternate source of air, and that the 
preheat supplied to the alternate air intake is not less than that 
provided by the engine cooling air downstream of the cylinders;
    (3) Each rotorcraft with altitude engines using conventional venturi 
carburetors has a preheater capable of providing a heat rise of 120 
degrees F.; and
    (4) Each rotorcraft with altitude engines using carburetors tending 
to prevent icing has a preheater that can provide a heat rise of--
    (i) 100 degrees F.; or
    (ii) If a fluid deicing system is used, at least 40 degrees F.
    (b) Turbine engine. (1) It must be shown that each turbine engine 
and its air inlet system can operate throughout the flight power range 
of the engine (including idling)--
    (i) Without accumulating ice on engine or inlet system components 
that would adversely affect engine operation or cause a serious loss of 
power under the icing conditions specified in appendix C of Part 29 of 
this chapter; and

[[Page 713]]

    (ii) In snow, both falling and blowing, without adverse effect on 
engine operation, within the limitations established for the rotorcraft.
    (2) Each turbine engine must idle for 30 minutes on the ground, with 
the air bleed available for engine icing protection at its critical 
condition, without adverse effect, in an atmosphere that is at a 
temperature between 15[deg] and 30 [deg]F (between -9[deg] and -1 
[deg]C) and has a liquid water content not less than 0.3 gram per cubic 
meter in the form of drops having a mean effective diameter not less 
than 20 microns, followed by momentary operation at takeoff power or 
thrust. During the 30 minutes of idle operation, the engine may be run 
up periodically to a moderate power or thrust setting in a manner 
acceptable to the Administrator.
    (c) Supercharged reciprocating engines. For each engine having 
superchargers to pressurize the air before it enters the carburetor, the 
heat rise in the air caused by that supercharging at any altitude may be 
utilized in determining compliance with paragraph (a) of this section if 
the heat rise utilized is that which will be available, automatically, 
for the applicable altitude and operating condition because of 
supercharging.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55470, Dec. 20, 1976; Amdt. 27-12, 42 FR 15045, Mar. 17, 1977; 
Amdt. 27-20, 49 FR 6849, Feb. 23, 1984; Amdt. 27-23, 53 FR 34214, Sept. 
2, 1988]

                             Exhaust System



Sec. 27.1121  General.

    For each exhaust system--
    (a) There must be means for thermal expansion of manifolds and 
pipes;
    (b) There must be means to prevent local hot spots;
    (c) Exhaust gases must discharge clear of the engine air intake, 
fuel system components, and drains;
    (d) Each exhaust system part with a surface hot enough to ignite 
flammable fluids or vapors must be located or shielded so that leakage 
from any system carrying flammable fluids or vapors will not result in a 
fire caused by impingement of the fluids or vapors on any part of the 
exhaust system including shields for the exhaust system;
    (e) Exhaust gases may not impair pilot vision at night due to glare;
    (f) If significant traps exist, each turbine engine exhaust system 
must have drains discharging clear of the rotorcraft, in any normal 
ground and flight attitudes, to prevent fuel accumulation after the 
failure of an attempted engine start;
    (g) Each exhaust heat exchanger must incorporate means to prevent 
blockage of the exhaust port after any internal heat exchanger failure.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964 as amended by Amdt. 27-12, 42 
FR 15045, Mar. 17, 1977]



Sec. 27.1123  Exhaust piping.

    (a) Exhaust piping must be heat and corrosion resistant, and must 
have provisions to prevent failure due to expansion by operating 
temperatures.
    (b) Exhaust piping must be supported to withstand any vibration and 
inertia loads to which it would be subjected in operations.
    (c) Exhaust piping connected to components between which relative 
motion could exist must have provisions for flexibility.

[Amdt. 27-11, 41 FR 55470, Dec. 20, 1976]

                   Powerplant Controls and Accessories



Sec. 27.1141  Powerplant controls: general.

    (a) Powerplant controls must be located and arranged under Sec. 
27.777 and marked under Sec. 27.1555.
    (b) Each flexible powerplant control must be approved.
    (c) Each control must be able to maintain any set position without--
    (1) Constant attention; or
    (2) Tendency to creep due to control loads or vibration.
    (d) Controls of powerplant valves required for safety must have--
    (1) For manual valves, positive stops or in the case of fuel valves 
suitable index provisions, in the open and closed position; and

[[Page 714]]

    (2) For power-assisted valves, a means to indicate to the flight 
crew when the valve--
    (i) Is in the fully open or fully closed position; or
    (ii) Is moving between the fully open and fully closed position.
    (e) For turbine engine powered rotorcraft, no single failure or 
malfunction, or probable combination thereof, in any powerplant control 
system may cause the failure of any powerplant function necessary for 
safety.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-12, 
42 FR 15045, Mar. 17, 1977; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988; 
Amdt. 27-33, 61 FR 21907, May 10, 1996]



Sec. 27.1143  Engine controls.

    (a) There must be a separate power control for each engine.
    (b) Power controls must be grouped and arranged to allow--
    (1) Separate control of each engine; and
    (2) Simultaneous control of all engines.
    (c) Each power control must provide a positive and immediately 
responsive means of controlling its engine.
    (d) If a power control incorporates a fuel shutoff feature, the 
control must have a means to prevent the inadvertent movement of the 
control into the shutoff position. The means must--
    (1) Have a positive lock or stop at the idle position; and
    (2) Require a separate and distinct operation to place the control 
in the shutoff position.
    (e) For rotorcraft to be certificated for a 30-second OEI power 
rating, a means must be provided to automatically activate and control 
the 30-second OEI power and prevent any engine from exceeding the 
installed engine limits associated with the 30-second OEI power rating 
approved for the rotorcraft.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55470, Dec. 20, 1976; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988; 
Amdt. 27-29, 59 FR 47767, Sept. 16, 1994]



Sec. 27.1145  Ignition switches.

    (a) There must be means to quickly shut off all ignition by the 
grouping of switches or by a master ignition control.
    (b) Each group of ignition switches, except ignition switches for 
turbine engines for which continuous ignition is not required, and each 
master ignition control must have a means to prevent its inadvertent 
operation.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-12, 
42 FR 15045, Mar. 17, 1977]



Sec. 27.1147  Mixture controls.

    If there are mixture controls, each engine must have a separate 
control and the controls must be arranged to allow--
    (a) Separate control of each engine; and
    (b) Simultaneous control of all engines.



Sec. 27.1151  Rotor brake controls.

    (a) It must be impossible to apply the rotor brake inadvertently in 
flight.
    (b) There must be means to warn the crew if the rotor brake has not 
been completely released before takeoff.

[Doc. No. 28008, 61 FR 21907, May 10, 1996]



Sec. 27.1163  Powerplant accessories.

    (a) Each engine-mounted accessory must--
    (1) Be approved for mounting on the engine involved;
    (2) Use the provisions on the engine for mounting; and
    (3) Be sealed in such a way as to prevent contamination of the 
engine oil system and the accessory system.
    (b) Unless other means are provided, torque limiting means must be 
provided for accessory drives located on any component of the 
transmission and rotor drive system to prevent damage to these 
components from excessive accessory load.

[Amdt. 27-2, 33 FR 964, Jan. 26, 1968, as amended by Amdt. 27-20, 49 FR 
6849, Feb. 23, 1984; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988]

[[Page 715]]

                       Powerplant Fire Protection



Sec. 27.1183  Lines, fittings, and components.

    (a) Except as provided in paragraph (b) of this section, each line, 
fitting, and other component carrying flammable fluid in any area 
subject to engine fire conditions must be fire resistant, except that 
flammable fluid tanks and supports which are part of and attached to the 
engine must be fireproof or be enclosed by a fireproof shield unless 
damage by fire to any non-fireproof part will not cause leakage or 
spillage of flammable fluid. Components must be shielded or located so 
as to safeguard against the ignition of leaking flammable fluid. An 
integral oil sump of less than 25-quart capacity on a reciprocating 
engine need not be fireproof nor be enclosed by a fireproof shield.
    (b) Paragraph (a) does not apply to--
    (1) Lines, fittings, and components which are already approved as 
part of a type certificated engine; and
    (2) Vent and drain lines, and their fittings, whose failure will not 
result in, or add to, a fire hazard.
    (c) Each flammable fluid drain and vent must discharge clear of the 
induction system air inlet.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-1, 32 
FR 6914, May 5, 1967; Amdt. 27-9, 39 FR 35462, Oct. 1, 1974; Amdt. 27-
20, 49 FR 6849, Feb. 23, 1984]



Sec. 27.1185  Flammable fluids.

    (a) Each fuel tank must be isolated from the engines by a firewall 
or shroud.
    (b) Each tank or reservoir, other than a fuel tank, that is part of 
a system containing flammable fluids or gases must be isolated from the 
engine by a firewall or shroud, unless the design of the system, the 
materials used in the tank and its supports, the shutoff means, and the 
connections, lines and controls provide a degree of safety equal to that 
which would exist if the tank or reservoir were isolated from the 
engines.
    (c) There must be at least one-half inch of clear airspace between 
each tank and each firewall or shroud isolating that tank, unless 
equivalent means are used to prevent heat transfer from each engine 
compartment to the flammable fluid.
    (d) Absorbent materials close to flammable fluid system components 
that might leak must be covered or treated to prevent the absorption of 
hazardous quantities of fluids.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 964, Jan. 26, 1968; Amdt. 27-11, 41 FR 55470, Dec. 20, 1976; Amdt. 
27-37, 64 FR 45095, Aug. 18, 1999]



Sec. 27.1187  Ventilation and drainage.

    Each compartment containing any part of the powerplant installation 
must have provision for ventilation and drainage of flammable fluids. 
The drainage means must be--
    (a) Effective under conditions expected to prevail when drainage is 
needed, and
    (b) Arranged so that no discharged fluid will cause an additional 
fire hazard.

[Doc. No. 29247, 64 FR 45095, Aug. 18, 1999]



Sec. 27.1189  Shutoff means.

    (a) There must be means to shut off each line carrying flammable 
fluids into the engine compartment, except--
    (1) Lines, fittings, and components forming an intergral part of an 
engine;
    (2) For oil systems for which all components of the system, 
including oil tanks, are fireproof or located in areas not subject to 
engine fire conditions; and
    (3) For reciprocating engine installations only, engine oil system 
lines in installation using engines of less than 500 cu. in. 
displacement.
    (b) There must be means to guard against inadvertent operation of 
each shutoff, and to make it possible for the crew to reopen it in 
flight after it has been closed.
    (c) Each shutoff valve and its control must be designed, located, 
and protected to function properly under any condition likely to result 
from an engine fire.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 964, Jan. 26, 1968; Amdt. 27-20, 49 FR 6850, Feb. 23, 1984; Amdt. 27-
23, 53 FR 34214, Sept. 2, 1988]

[[Page 716]]



Sec. 27.1191  Firewalls.

    (a) Each engine, including the combustor, turbine, and tailpipe 
sections of turbine engines must be isolated by a firewall, shroud, or 
equivalent means, from personnel compartments, structures, controls, 
rotor mechanisms, and other parts that are--
    (1) Essential to a controlled landing: and
    (2) Not protected under Sec. 27.861.
    (b) Each auxiliary power unit and combustion heater, and any other 
combustion equipment to be used in flight, must be isolated from the 
rest of the rotorcraft by firewalls, shrouds, or equivalent means.
    (c) In meeting paragraphs (a) and (b) of this section, account must 
be taken of the probable path of a fire as affected by the airflow in 
normal flight and in autorotation.
    (d) Each firewall and shroud must be constructed so that no 
hazardous quantity of air, fluids, or flame can pass from any engine 
compartment to other parts of the rotorcraft.
    (e) Each opening in the firewall or shroud must be sealed with 
close-fitting, fireproof grommets, bushings, or firewall fittings.
    (f) Each firewall and shroud must be fireproof and protected against 
corrosion.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 22 
FR 964, Jan. 26, 1968]



Sec. 27.1193  Cowling and engine compartment covering.

    (a) Each cowling and engine compartment covering must be constructed 
and supported so that it can resist the vibration, inertia, and air 
loads to which it may be subjected in operation.
    (b) There must be means for rapid and complete drainage of each part 
of the cowling or engine compartment in the normal ground and flight 
attitudes.
    (c) No drain may discharge where it might cause a fire hazard.
    (d) Each cowling and engine compartment covering must be at least 
fire resistant.
    (e) Each part of the cowling or engine compartment covering subject 
to high temperatures due to its nearness to exhaust system parts or 
exhaust gas impingement must be fireproof.
    (f) A means of retaining each openable or readily removable panel, 
cowling, or engine or rotor drive system covering must be provided to 
preclude hazardous damage to rotors or critical control components in 
the event of structural or mechanical failure of the normal retention 
means, unless such failure is extremely improbable.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 
53 FR 34214, Sept. 2, 1988]



Sec. 27.1194  Other surfaces.

    All surfaces aft of, and near, powerplant compartments, other than 
tail surfaces not subject to heat, flames, or sparks emanating from a 
powerplant compartment, must be at least fire resistant.

[Amdt. 27-2, 33 FR 964, Jan. 26, 1968]



Sec. 27.1195  Fire detector systems.

    Each turbine engine powered rotorcraft must have approved quick-
acting fire detectors in numbers and locations insuring prompt detection 
of fire in the engine compartment which cannot be readily observed in 
flight by the pilot in the cockpit.

[Amdt. 27-5, 36 FR 5493, Mar. 24, 1971]



                           Subpart F_Equipment

                                 General



Sec. 27.1301  Function and installation.

    Each item of installed equipment must--
    (a) Be of a kind and design appropriate to its intended function;
    (b) Be labeled as to its identification, function, or operating 
limitations, or any applicable combination of these factors;
    (c) Be installed according to limitations specified for that 
equipment; and
    (d) Function properly when installed.



Sec. 27.1303  Flight and navigation instruments.

    The following are the required flight and navigation instruments:
    (a) An airspeed indicator.
    (b) An altimeter.
    (c) A magnetic direction indicator.

[[Page 717]]



Sec. 27.1305  Powerplant instruments.

    The following are the required powerplant instruments:
    (a) A carburetor air temperature indicator, for each engine having a 
preheater that can provide a heat rise in excess of 60 [deg]F.
    (b) A cylinder head temperature indicator, for each--
    (1) Air cooled engine;
    (2) Rotorcraft with cooling shutters; and
    (3) Rotorcraft for which compliance with Sec. 27.1043 is shown in 
any condition other than the most critical flight condition with respect 
to cooling.
    (c) A fuel pressure indicator, for each pump-fed engine.
    (d) A fuel quantity indicator, for each fuel tank.
    (e) A manifold pressure indicator, for each altitude engine.
    (f) An oil temperature warning device to indicate when the 
temperature exceeds a safe value in each main rotor drive gearbox 
(including any gearboxes essential to rotor phasing) having an oil 
system independent of the engine oil system.
    (g) An oil pressure warning device to indicate when the pressure 
falls below a safe value in each pressure-lubricated main rotor drive 
gearbox (including any gearboxes essential to rotor phasing) having an 
oil system independent of the engine oil system.
    (h) An oil pressure indicator for each engine.
    (i) An oil quantity indicator for each oil tank.
    (j) An oil temperature indicator for each engine.
    (k) At least one tachometer to indicate the r.p.m. of each engine 
and, as applicable--
    (1) The r.p.m. of the single main rotor;
    (2) The common r.p.m. of any main rotors whose speeds cannot vary 
appreciably with respect to each other; or
    (3) The r.p.m. of each main rotor whose speed can vary appreciably 
with respect to that of another main rotor.
    (l) A low fuel warning device for each fuel tank which feeds an 
engine. This device must--
    (1) Provide a warning to the flightcrew when approximately 10 
minutes of usable fuel remains in the tank; and
    (2) Be independent of the normal fuel quantity indicating system.
    (m) Means to indicate to the flightcrew the failure of any fuel pump 
installed to show compliance with Sec. 27.955.
    (n) A gas temperature indicator for each turbine engine.
    (o) Means to enable the pilot to determine the torque of each 
turboshaft engine, if a torque limitation is established for that engine 
under Sec. 27.1521(e).
    (p) For each turbine engine, an indicator to indicate the 
functioning of the powerplant ice protection system.
    (q) An indicator for the fuel filter required by Sec. 27.997 to 
indicate the occurrence of contamination of the filter at the degree 
established by the applicant in compliance with Sec. 27.955.
    (r) For each turbine engine, a warning means for the oil strainer or 
filter required by Sec. 27.1019, if it has no bypass, to warn the pilot 
of the occurrence of contamination of the strainer or filter before it 
reaches the capacity established in accordance with Sec. 27.1019(a)(2).
    (s) An indicator to indicate the functioning of any selectable or 
controllable heater used to prevent ice clogging of fuel system 
components.
    (t) For rotorcraft for which a 30-second/2-minute OEI power rating 
is requested, a means must be provided to alert the pilot when the 
engine is at the 30-second and the 2-minute OEI power levels, when the 
event begins, and when the time interval expires.
    (u) For each turbine engine utilizing 30-second/2-minute OEI power, 
a device or system must be provided for use by ground personnel which--
    (1) Automatically records each usage and duration of power at the 
30-second and 2-minute OEI levels;
    (2) Permits retrieval of the recorded data;
    (3) Can be reset only by ground maintenance personnel; and
    (4) Has a means to verify proper operation of the system or device.

[[Page 718]]

    (v) Warning or caution devices to signal to the flight crew when 
ferromagnetic particles are detected by the chip detector required by 
Sec. 27.1337(e).

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-9, 39 
FR 35462, Oct. 1, 1974; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988; Amdt. 
27-29, 59 FR 47767, Sept. 16, 1994; Amdt. 27-37, 64 FR 45095, Aug. 18, 
1999; 64 FR 47563, Aug. 31, 1999]



Sec. 27.1307  Miscellaneous equipment.

    The following is the required miscellaneous equipment:
    (a) An approved seat for each occupant.
    (b) An approved safety belt for each occupant.
    (c) A master switch arrangement.
    (d) An adequate source of electrical energy, where electrical energy 
is necessary for operation of the rotorcraft.
    (e) Electrical protective devices.



Sec. 27.1309  Equipment, systems, and installations.

    (a) The equipment, systems, and installations whose functioning is 
required by this subchapter must be designed and installed to ensure 
that they perform their intended functions under any foreseeable 
operating condition.
    (b) The equipment, systems, and installations of a multiengine 
rotorcraft must be designed to prevent hazards to the rotorcraft in the 
event of a probable malfunction or failure.
    (c) The equipment, systems, and installations of single-engine 
rotorcraft must be designed to minimize hazards to the rotorcraft in the 
event of a probable malfunction or failure.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 
49 FR 44435, Nov. 6, 1984; Amdt. 27-46, 76 FR 33135, June 8, 2011]



Sec. 27.1316  Electrical and electronic system lightning protection.

    (a) Each electrical and electronic system that performs a function, 
for which failure would prevent the continued safe flight and landing of 
the rotorcraft, must be designed and installed so that--
    (1) The function is not adversely affected during and after the time 
the rotorcraft is exposed to lightning; and
    (2) The system automatically recovers normal operation of that 
function in a timely manner after the rotorcraft is exposed to 
lightning.
    (b) For rotorcraft approved for instrument flight rules operation, 
each electrical and electronic system that performs a function, for 
which failure would reduce the capability of the rotorcraft or the 
ability of the flightcrew to respond to an adverse operating condition, 
must be designed and installed so that the function recovers normal 
operation in a timely manner after the rotorcraft is exposed to 
lightning.

[Doc. No. FAA-2010-0224, Amdt. 27-46, 76 FR 33135, June 8, 2011]



Sec. 27.1317  High-intensity Radiated Fields (HIRF) Protection.

    (a) Except as provided in paragraph (d) of this section, each 
electrical and electronic system that performs a function whose failure 
would prevent the continued safe flight and landing of the rotorcraft 
must be designed and installed so that--
    (1) The function is not adversely affected during and after the time 
the rotorcraft is exposed to HIRF environment I, as described in 
appendix D to this part;
    (2) The system automatically recovers normal operation of that 
function, in a timely manner, after the rotorcraft is exposed to HIRF 
environment I, as described in appendix D to this part, unless this 
conflicts with other operational or functional requirements of that 
system;
    (3) The system is not adversely affected during and after the time 
the rotorcraft is exposed to HIRF environment II, as described in 
appendix D to this part; and
    (4) Each function required during operation under visual flight 
rules is not adversely affected during and after the time the rotorcraft 
is exposed to HIRF environment III, as described in appendix D to this 
part.
    (b) Each electrical and electronic system that performs a function 
whose failure would significantly reduce the capability of the 
rotorcraft or the ability of the flightcrew to respond to an adverse 
operating condition must be designed and installed so the system is

[[Page 719]]

not adversely affected when the equipment providing these functions is 
exposed to equipment HIRF test level 1 or 2, as described in appendix D 
to this part.
    (c) Each electrical and electronic system that performs a function 
whose failure would reduce the capability of the rotorcraft or the 
ability of the flightcrew to respond to an adverse operating condition, 
must be designed and installed so the system is not adversely affected 
when the equipment providing these functions is exposed to equipment 
HIRF test level 3, as described in appendix D to this part.
    (d) Before December 1, 2012, an electrical or electronic system that 
performs a function whose failure would prevent the continued safe 
flight and landing of a rotorcraft may be designed and installed without 
meeting the provisions of paragraph (a) provided--
    (1) The system has previously been shown to comply with special 
conditions for HIRF, prescribed under Sec. 21.16, issued before 
December 1, 2007;
    (2) The HIRF immunity characteristics of the system have not changed 
since compliance with the special conditions was demonstrated; and
    (3) The data used to demonstrate compliance with the special 
conditions is provided.

[Doc. No. FAA-2006-23657, 72 FR 44026, Aug. 6, 2007]

                        Instruments: Installation



Sec. 27.1321  Arrangement and visibility.

    (a) Each flight, navigation, and powerplant instrument for use by 
any pilot must be easily visible to him.
    (b) For each multiengine rotorcraft, identical powerplant 
instruments must be located so as to prevent confusion as to which 
engine each instrument relates.
    (c) Instrument panel vibration may not damage, or impair the 
readability or accuracy of, any instrument.
    (d) If a visual indicator is provided to indicate malfunction of an 
instrument, it must be effective under all probable cockpit lighting 
conditions.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964; 29 FR 17885, Dec. 17, 1964, 
as amended by Amdt. 27-13, 42 FR 36971, July 18, 1977]



Sec. 27.1322  Warning, caution, and advisory lights.

    If warning, caution or advisory lights are installed in the cockpit, 
they must, unless otherwise approved by the Administrator, be--
    (a) Red, for warning lights (lights indicating a hazard which may 
require immediate corrective action):
    (b) Amber, for caution lights (lights indicating the possible need 
for future corrective action);
    (c) Green, for safe operation lights; and
    (d) Any other color, including white, for lights not described in 
paragraphs (a) through (c) of this section, provided the color differs 
sufficiently from the colors prescribed in paragraphs (a) through (c) of 
this section to avoid possible confusion.

[Amdt. 27-11, 41 FR 55470, Dec. 20, 1976]



Sec. 27.1323  Airspeed indicating system.

    (a) Each airspeed indicating instrument must be calibrated to 
indicate true airspeed (at sea level with a standard atmosphere) with a 
minimum practicable instrument calibration error when the corresponding 
pitot and static pressures are applied.
    (b) The airspeed indicating system must be calibrated in flight at 
forward speeds of 20 knots and over.
    (c) At each forward speed above 80 percent of the climbout speed, 
the airspeed indicator must indicate true airspeed, at sea level with a 
standard atmosphere, to within an allowable installation error of not 
more than the greater of--
    (1) 3 percent of the calibrated airspeed; or

[[Page 720]]

    (2) Five knots.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-13, 
42 FR 36972, July 18, 1977]



Sec. 27.1325  Static pressure systems.

    (a) Each instrument with static air case connections must be vented 
so that the influence of rotorcraft speed, the opening and closing of 
windows, airflow variation, and moisture or other foreign matter does 
not seriously affect its accuracy.
    (b) Each static pressure port must be designed and located in such 
manner that the correlation between air pressure in the static pressure 
system and true ambient atmospheric static pressure is not altered when 
the rotorcraft encounters icing conditions. An anti-icing means or an 
alternate source of static pressure may be used in showing compliance 
with this requirement. If the reading of the altimeter, when on the 
alternate static pressure system, differs from the reading of the 
altimeter when on the primary static system by more than 50 feet, a 
correction card must be provided for the alternate static system.
    (c) Except as provided in paragraph (d) of this section, if the 
static pressure system incorporates both a primary and an alternate 
static pressure source, the means for selecting one or the other source 
must be designed so that--
    (1) When either source is selected, the other is blocked off; and
    (2) Both sources cannot be blocked off simultaneously.
    (d) For unpressurized rotorcraft, paragraph (c)(1) of this section 
does not apply if it can be demonstrated that the static pressure system 
calibration, when either static pressure source is selected is not 
changed by the other static pressure source being open or blocked.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-13, 
42 FR 36972, July 18, 1977]



Sec. 27.1327  Magnetic direction indicator.

    (a) Except as provided in paragraph (b) of this section--
    (1) Each magnetic direction indicator must be installed so that its 
accuracy is not excessively affected by the rotorcraft's vibration or 
magnetic fields; and
    (2) The compensated installation may not have a deviation, in level 
flight, greater than 10 degrees on any heading.
    (b) A magnetic nonstabilized direction indicator may deviate more 
than 10 degrees due to the operation of electrically powered systems 
such as electrically heated windshields if either a magnetic stabilized 
direction indicator, which does not have a deviation in level flight 
greater than 10 degrees on any heading, or a gyroscopic direction 
indicator, is installed. Deviations of a magnetic nonstabilized 
direction indicator of more than 10 degrees must be placarded in 
accordance with Sec. 27.1547(e).

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 27-13, 42 FR 36972, July 18, 1977]



Sec. 27.1329  Automatic pilot system.

    (a) Each automatic pilot system must be designed so that the 
automatic pilot can--
    (1) Be sufficiently overpowered by one pilot to allow control of the 
rotorcraft; and
    (2) Be readily and positively disengaged by each pilot to prevent it 
from interfering with control of the rotorcraft.
    (b) Unless there is automatic synchronization, each system must have 
a means to readily indicate to the pilot the alignment of the actuating 
device

[[Page 721]]

in relation to the control system it operates.
    (c) Each manually operated control for the system's operation must 
be readily accessible to the pilots.
    (d) The system must be designed and adjusted so that, within the 
range of adjustment available to the pilot, it cannot produce hazardous 
loads on the rotorcraft or create hazardous deviations in the flight 
path under any flight condition appropriate to its use, either during 
normal operation or in the event of a malfunction, assuming that 
corrective action begins within a reasonable period of time.
    (e) If the automatic pilot integrates signals from auxiliary 
controls or furnishes signals for operation of other equipment, there 
must be positive interlocks and sequencing of engagement to prevent 
improper operation.
    (f) If the automatic pilot system can be coupled to airborne 
navigation equipment, means must be provided to indicate to the pilots 
the current mode of operation. Selector switch position is not 
acceptable as a means of indication.

[Amdt. 27-21, 49 FR 44435, Nov. 6, 1984, as amended by Amdt. 27-35, 63 
FR 43285, Aug. 12, 1998]



Sec. 27.1335  Flight director systems.

    If a flight director system is installed, means must be provided to 
indicate to the flight crew its current mode of operation. Selector 
switch position is not acceptable as a means of indication.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 27-13, 42 FR 36972, July 18, 1977]



Sec. 27.1337  Powerplant instruments.

    (a) Instruments and instrument lines. (1) Each powerplant instrument 
line must meet the requirements of Sec. Sec. 27.- 961 and 27.993.
    (2) Each line carrying flammable fluids under pressure must--
    (i) Have restricting orifices or other safety devices at the source 
of pressure to prevent the escape of excessive fluid if the line fails; 
and
    (ii) Be installed and located so that the escape of fluids would not 
create a hazard.
    (3) Each powerplant instrument that utilizes flammable fluids must 
be installed and located so that the escape of fluid would not create a 
hazard.
    (b) Fuel quantity indicator. Each fuel quantity indicator must be 
installed to clearly indicate to the flight crew the quantity of fuel in 
each tank in flight. In addition--
    (1) Each fuel quantity indicator must be calibrated to read ``zero'' 
during level flight when the quantity of fuel remaining in the tank is 
equal to the unusable fuel supply determined under Sec. 27.959;
    (2) When two or more tanks are closely interconnected by a gravity 
feed system and vented, and when it is impossible to feed from each tank 
separately, at least one fuel quantity indicator must be installed; and
    (3) Each exposed sight gauge used as a fuel quantity indicator must 
be protected against damage.
    (c) Fuel flowmeter system. If a fuel flowmeter system is installed, 
each metering component must have a means for bypassing the fuel supply 
if malfunction of that component severely restricts fuel flow.
    (d) Oil quantity indicator. There must be means to indicate the 
quantity of oil in each tank--
    (1) On the ground (including during the filling of each tank); and
    (2) In flight, if there is an oil transfer system or reserve oil 
supply system.
    (e) Rotor drive system transmissions and gearboxes utilizing 
ferromagnetic materials must be equipped with chip detectors designed to 
indicate the presence of ferromagnetic particles resulting from damage 
or excessive wear. Chip detectors must--
    (1) Be designed to provide a signal to the device required by Sec. 
27.1305(v) and be provided with a means to allow crewmembers to check, 
in flight, the function of each detector electrical circuit and signal.

[[Page 722]]

    (2) [Reserved]

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c) 49 U.S.C. 1655(c))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-12, 
42 FR 15046, Mar. 17, 1977; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988; 
Amdt. 27-37, 64 FR 45095, Aug. 18, 1999]

                    Electrical Systems and Equipment



Sec. 27.1351  General.

    (a) Electrical system capacity. Electrical equipment must be 
adequate for its intended use. In addition--
    (1) Electric power sources, their transmission cables, and their 
associated control and protective devices must be able to furnish the 
required power at the proper voltage to each load circuit essential for 
safe operation; and
    (2) Compliance with paragraph (a)(1) of this section must be shown 
by an electrical load analysis, or by electrical measurements that take 
into account the electrical loads applied to the electrical system, in 
probable combinations and for probable durations.
    (b) Function. For each electrical system, the following apply:
    (1) Each system, when installed, must be--
    (i) Free from hazards in itself, in its method of operation, and in 
its effects on other parts of the rotorcraft; and
    (ii) Protected from fuel, oil, water, other detrimental substances, 
and mechanical damage.
    (2) Electric power sources must function properly when connected in 
combination or independently.
    (3) No failure or malfunction of any source may impair the ability 
of any remaining source to supply load circuits essential for safe 
operation.
    (4) Each electric power source control must allow the independent 
operation of each source.
    (c) Generating system. There must be at least one generator if the 
system supplies power to load circuits essential for safe operation. In 
addition--
    (1) Each generator must be able to deliver its continuous rated 
power;
    (2) Generator voltage control equipment must be able to dependably 
regulate each generator output within rated limits;
    (3) Each generator must have a reverse current cutout designed to 
disconnect the generator from the battery and from the other generators 
when enough reverse current exists to damage that generator; and
    (4) Each generator must have an overvoltage control designed and 
installed to prevent damage to the electrical system, or to equipment 
supplied by the electrical system, that could result if that generator 
were to develop an overvoltage condition.
    (d) Instruments. There must be means to indicate to appropriate 
crewmembers the electric power system quantities essential for safe 
operation of the system. In addition--
    (1) For direct current systems, an ammeter that can be switched into 
each generator feeder may be used; and
    (2) If there is only one generator, the ammeter may be in the 
battery feeder.
    (e) External power. If provisions are made for connecting external 
power to the rotorcraft, and that external power can be electrically 
connected to equipment other than that used for engine starting, means 
must be provided to ensure that no external power supply having a 
reverse polarity, or a reverse phase sequence, can supply power to the 
rotorcraft's electrical system.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55470, Dec. 20, 1976; Amdt. 27-13, 42 FR 36972, July 18, 1977]



Sec. 27.1353  Storage battery design and installation.

    (a) Each storage battery must be designed and installed as 
prescribed in this section.
    (b) Safe cell temperatures and pressures must be maintained during 
any probable charging and discharging condition. No uncontrolled 
increase in cell temperature may result when the battery is recharged 
(after previous complete discharge)--
    (1) At maximum regulated voltage or power;
    (2) During a flight of maximum duration; and

[[Page 723]]

    (3) Under the most adverse cooling condition likely to occur in 
service.
    (c) Compliance with paragraph (b) of this section must be shown by 
test unless experience with similar batteries and installations has 
shown that maintaining safe cell temperatures and pressures presents no 
problem.
    (d) No explosive or toxic gases emitted by any battery in normal 
operation, or as the result of any probable malfunction in the charging 
system or battery installation, may accumulate in hazardous quantities 
within the rotorcraft.
    (e) No corrosive fluids or gases that may escape from the battery 
may damage surrounding structures or adjacent essential equipment.
    (f) Each nickel cadmium battery installation capable of being used 
to start an engine or auxiliary power unit must have provisions to 
prevent any hazardous effect on structure or essential systems that may 
be caused by the maximum amount of heat the battery can generate during 
a short circuit of the battery or of its individual cells.
    (g) Nickel cadmium battery installations capable of being used to 
start an engine or auxiliary power unit must have--
    (1) A system to control the charging rate of the battery 
automatically so as to prevent battery overheating;
    (2) A battery temperature sensing and over-temperature warning 
system with a means for disconnecting the battery from its charging 
source in the event of an over-temperature condition; or
    (3) A battery failure sensing and warning system with a means for 
disconnecting the battery from its charging source in the event of 
battery failure.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-13, 
42 FR 36972, July 18, 1977; Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]



Sec. 27.1357  Circuit protective devices.

    (a) Protective devices, such as fuses or circuit breakers, must be 
installed in each electrical circuit other than--
    (1) The main circuits of starter motors; and
    (2) Circuits in which no hazard is presented by their omission.
    (b) A protective device for a circuit essential to flight safety may 
not be used to protect any other circuit.
    (c) Each resettable circuit protective device (``trip free'' device 
in which the tripping mechanism cannot be overridden by the operating 
control) must be designed so that--
    (1) A manual operation is required to restore service after 
trippling; and
    (2) If an overload or circuit fault exists, the device will open the 
circuit regardless of the position of the operating control.
    (d) If the ability to reset a circuit breaker or replace a fuse is 
essential to safety in flight, that circuit breaker or fuse must be 
located and identified so that it can be readily reset or replaced in 
flight.
    (e) If fuses are used, there must be one spare of each rating, or 50 
percent spare fuses of each rating, whichever is greater.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964; 29 FR 17885, Dec. 17, 1964, 
as amended by Amdt. 27-13, 42 FR 36972, July 18, 1977]



Sec. 27.1361  Master switch.

    (a) There must be a master switch arrangement to allow ready 
disconnection of each electric power source from the main bus. The point 
of disconnection must be adjacent to the sources controlled by the 
switch.
    (b) Load circuits may be connected so that they remain energized 
after the switch is opened, if they are protected by circuit protective 
devices, rated at five amperes or less, adjacent to the electric power 
source.
    (c) The master switch or its controls must be installed so that the 
switch is easily discernible and accessible to a crewmember in flight.



Sec. 27.1365  Electric cables.

    (a) Each electric connecting cable must be of adequate capacity.

[[Page 724]]

    (b) Each cable that would overheat in the event of circuit overload 
or fault must be at least flame resistant and may not emit dangerous 
quantities of toxic fumes.
    (c) Insulation on electrical wire and cable installed in the 
rotorcraft must be self-extinguishing when tested in accordance with 
appendix F, part I(a)(3), of part 25 of this chapter.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-35, 
63 FR 43285, Aug. 12, 1998]



Sec. 27.1367  Switches.

    Each switch must be--
    (a) Able to carry its rated current;
    (b) Accessible to the crew; and
    (c) Labeled as to operation and the circuit controlled.

                                 Lights



Sec. 27.1381  Instrument lights.

    The instrument lights must--
    (a) Make each instrument, switch, and other devices for which they 
are provided easily readable; and
    (b) Be installed so that--
    (1) Their direct rays are shielded from the pilot's eyes; and
    (2) No objectionable reflections are visible to the pilot.



Sec. 27.1383  Landing lights.

    (a) Each required landing or hovering light must be approved.
    (b) Each landing light must be installed so that--
    (1) No objectionable glare is visible to the pilot;
    (2) The pilot is not adversely affected by halation; and
    (3) It provides enough light for night operation, including hovering 
and landing.
    (c) At least one separate switch must be provided, as applicable--
    (1) For each separately installed landing light; and
    (2) For each group of landing lights installed at a common location.



Sec. 27.1385  Position light system installation.

    (a) General. Each part of each position light system must meet the 
applicable requirements of this section, and each system as a whole must 
meet the requirements of Sec. Sec. 27.1387 through 27.1397.
    (b) Forward position lights. Forward position lights must consist of 
a red and a green light spaced laterally as far apart as practicable and 
installed forward on the rotorcraft so that, with the rotorcraft in the 
normal flying position, the red light is on the left side and the green 
light is on the right side. Each light must be approved.
    (c) Rear position light. The rear position light must be a white 
light mounted as far aft as practicable, and must be approved.
    (d) Circuit. The two forward position lights and the rear position 
light must make a single circuit.
    (e) Light covers and color filters. Each light cover or color filter 
must be at least flame resistant and may not change color or shape or 
lose any appreciable light transmission during normal use.



Sec. 27.1387  Position light system dihedral angles.

    (a) Except as provided in paragraph (e) of this section, each 
forward and rear position light must, as installed, show unbroken light 
within the dihedral angles described in this section.
    (b) Dihedral angle L (left) is formed by two intersecting vertical 
planes, the first parallel to the longitudinal axis of the rotorcraft, 
and the other at 110 degrees to the left of the first, as viewed when 
looking forward along the longitudinal axis.
    (c) Dihedral angle R (right) is formed by two intersecting vertical 
planes, the first parallel to the longitudinal axis of the rotorcraft, 
and the other at 110 degrees to the right of the first, as viewed when 
looking forward along the longitudinal axis.
    (d) Dihedral angle A (aft) is formed by two intersecting vertical 
planes making angles of 70 degrees to the right and to the left, 
respectively, to a vertical plane passing through the longitudinal axis, 
as viewed when looking aft along the longitudinal axis.
    (e) If the rear position light, when mounted as far aft as 
practicable in accordance with Sec. 25.1385(c), cannot show unbroken 
light within dihedral angle A

[[Page 725]]

(as defined in paragraph (d) of this section), a solid angle or angles 
of obstructed visibility totaling not more than 0.04 steradians is 
allowable within that dihedral angle, if such solid angle is within a 
cone whose apex is at the rear position light and whose elements make an 
angle of 30[deg] with a vertical line passing through the rear position 
light.

(49 U.S.C. 1655(c))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-7, 36 
FR 21278, Nov. 5, 1971]



Sec. 27.1389  Position light distribution and intensities.

    (a) General. the intensities prescribed in this section must be 
provided by new equipment with light covers and color filters in place. 
Intensities must be determined with the light source operating at a 
steady value equal to the average luminous output of the source at the 
normal operating voltage of the rotorcraft. The light distribution and 
intensity of each position light must meet the requirements of paragraph 
(b) of this section.
    (b) Forward and rear position lights. The light distribution and 
intensities of forward and rear position lights must be expressed in 
terms of minimum intensities in the horizontal plane, minimum 
intensities in any vertical plane, and maximum intensities in 
overlapping beams, within dihedral angles L, R, and A, and must meet the 
following requirements:
    (1) Intensities in the horizontal plane. Each intensity in the 
horizontal plane (the plane containing the longitudinal axis of the 
rotorcraft and perpendicular to the plane of symmetry of the rotorcraft) 
must equal or exceed the values in Sec. 27.1391.
    (2) Intensities in any vertical plane. Each intensity in any 
vertical plane (the plane perpendicular to the horizontal plane) must 
equal or exceed the appropriate value in Sec. 27.1393, where I is the 
minimum intensity prescribed in Sec. 27.1391 for the corresponding 
angles in the horizontal plane.
    (3) Intensities in overlaps between adjacent signals. No intensity 
in any overlap between adjacent signals may exceed the values in Sec. 
27.1395, except that higher intensities in overlaps may be used with 
main beam intensities substantially greater than the minima specified in 
Sec. Sec. 27.1391 and 27.1393, if the overlap intensities in relation 
to the main beam intensities do not adversely affect signal clarity. 
When the peak intensity of the forward position lights is greater than 
100 candles, the maximum overlap intensities between them may exceed the 
values in Sec. 27.1395 if the overlap intensity in Area A is not more 
than 10 percent of peak position light intensity and the overlap 
intensity in Area B is not more than 2.5 percent of peak position light 
intensity.



Sec. 27.1391  Minimum intensities in the horizontal plane of forward 
and rear position lights.

    Each position light intensity must equal or exceed the applicable 
values in the following table:

------------------------------------------------------------------------
                                        Angle from right or
                                       left of longitudinal    Intensity
   Dihedral angle (light included)      axis, measured from    (candles)
                                            dead ahead
------------------------------------------------------------------------
L and R (forward red and green).....  10[deg] to 10[deg]....          40
                                      10[deg] to 20[deg]....          30
                                      20[deg] to 110[deg]...           5
A (rear white)......................  110[deg] to 180[deg]..          20
------------------------------------------------------------------------



Sec. 27.1393  Minimum intensities in any vertical plane of forward
and rear position lights.

    Each position light intensity must equal or exceed the applicable 
values in the following table:

------------------------------------------------------------------------
                                                              Intensity,
          Angle above or below the horizontal plane                l
------------------------------------------------------------------------
0[deg]......................................................        1.00
0[deg] to 5[deg]............................................        0.90
5[deg] to 10[deg]...........................................        0.80
10[deg] to 15[deg]..........................................        0.70
15[deg] to 20[deg]..........................................        0.50
20[deg] to 30[deg]..........................................        0.30
30[deg] to 40[deg]..........................................        0.10
40[deg] to 90[deg]..........................................        0.05
------------------------------------------------------------------------



Sec. 27.1395  Maximum intensities in overlapping beams of forward and
rear position lights.

    No position light intensity may exceed the applicable values in the 
following table, except as provided in Sec. 27.1389(b)(3).

------------------------------------------------------------------------
                                                     Maximum Intensity
                                                 -----------------------
                    Overlaps                        Area A      Area B
                                                   (candles)   (candles)
------------------------------------------------------------------------
Green in dihedral angle L.......................          10           1
Red in dihedral angle R.........................          10           1

[[Page 726]]

 
Green in dihedral angle A.......................           5           1
Red in dihedral angle A.........................           5           1
Rear white in dihedral angle L..................           5           1
Rear white in dihedral angle R..................           5           1
------------------------------------------------------------------------


Where--
    (a) Area A includes all directions in the adjacent dihedral angle 
that pass through the light source and intersect the common boundary 
plane at more than 10 degrees but less than 20 degrees, and
    (b) Area B includes all directions in the adjacent dihedral angle 
that pass through the light source and intersect the common boundary 
plane at more than 20 degrees.



Sec. 27.1397  Color specifications.

    Each position light color must have the applicable International 
Commission on Illumination chromaticity coordinates as follows:
    (a) Aviation red--

    y is not greater than 0.335; and
    z is not greater than 0.002.

    (b) Aviation green--

    x is not greater than 0.440-0.320y;
    x is not greater than y-0.170; and
    y is not less than 0.390-0.170x.

    (c) Aviation white--

    x is not less than 0.300 and not greater than 0.540;
    y is not less than x-0.040'' or yc-0.010, whichever is 
the smaller; and
    y is not greater than x+0.020 nor 0.636-0.400x;
    Where yc is the y coordinate of the Planckian radiator 
for the value of x considered.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-6, 36 
FR 12972, July 10, 1971]



Sec. 27.1399  Riding light.

    (a) Each riding light required for water operation must be installed 
so that it can--
    (1) Show a white light for at least two nautical miles at night 
under clear atmospheric conditions; and
    (2) Show a maximum practicable unbroken light with the rotorcraft on 
the water.
    (b) Externally hung lights may be used.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 964, Jan. 26, 1968]



Sec. 27.1401  Anticollision light system.

    (a) General. If certification for night operation is requested, the 
rotorcraft must have an anticollision light system that--
    (1) Consists of one or more approved anticollision lights located so 
that their emitted light will not impair the crew's vision or detract 
from the conspicuity of the position lights; and
    (2) Meets the requirements of paragraphs (b) through (f) of this 
section.
    (b) Field of coverage. The system must consist of enough lights to 
illuminate the vital areas around the rotorcraft, considering the 
physical configuration and flight characteristics of the rotorcraft. The 
field of coverage must extend in each direction within at least 30 
degrees below the horizontal plane of the rotorcraft, except that there 
may be solid angles of obstructed visibility totaling not more than 0.5 
steradians.
    (c) Flashing characteristics. The arrangement of the system, that 
is, the number of light sources, beam width, speed of rotation, and 
other characteristics, must give an effective flash frequency of not 
less than 40, nor more than 100, cycles per minute. The effective flash 
frequency is the frequency at which the rotorcraft's complete 
anticollision light system is observed from a distance, and applies to 
each sector of light including any overlaps that exist when the system 
consists of more than one light source. In overlaps, flash frequencies 
may exceed 100, but not 180, cycles per minute.
    (d) Color. Each anticollision light must be aviation red and must 
meet the applicable requirements of Sec. 27.1397.
    (e) Light intensity. The minimum light intensities in any vertical 
plane, measured with the red filter (if used) and expressed in terms of 
``effective'' intensities, must meet the requirements of paragraph (f) 
of this section. The following relation must be assumed:

[[Page 727]]

[GRAPHIC] [TIFF OMITTED] TC28SE91.086

where:

Ie=effective intensity (candles).
I(t)=instantaneous intensity as a function of time.
t2-t1=flash time interval (seconds).
Normally, the maximum value of effective intensity is obtained when 
t2 and t1 are chosen so that the effective 
intensity is equal to the instantaneous intensity at t2 and 
t1.

    (f) Minimum effective intensities for anticollision light. Each 
anticollision light effective intensity must equal or exceed the 
applicable values in the following table:

------------------------------------------------------------------------
                                                               Effective
          Angle above or below the horizontal plane            intensity
                                                               (candles)
------------------------------------------------------------------------
0[deg] to 5[deg]............................................         150
5[deg] to 10[deg]...........................................          90
10[deg] to 20[deg]..........................................          30
20[deg] to 30[deg]..........................................          15
------------------------------------------------------------------------


[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-6, 36 
FR 12972, July 10, 1971; Amdt. 27-10, 41 FR 5290, Feb. 5, 1976]

                            Safety Equipment



Sec. 27.1411  General.

    (a) Required safety equipment to be used by the crew in an 
emergency, such as flares and automatic liferaft releases, must be 
readily accessible.
    (b) Stowage provisions for required safety equipment must be 
furnished and must--
    (1) Be arranged so that the equipment is directly accessible and its 
location is obvious; and
    (2) Protect the safety equipment from damage caused by being 
subjected to the inertia loads specified in Sec. 27.561.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55470, Dec. 20, 1976]



Sec. 27.1413  Safety belts.

    Each safety belt must be equipped with a metal to metal latching 
device.

(Secs. 313, 314, and 601 through 610 of the Federal Aviation Act of 1958 
(49 U.S.C. 1354, 1355, and 1421 through 1430) and sec. 6(c), Dept. of 
Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-15, 
43 FR 46233, Oct. 5, 1978; Amdt. 27-21, 49 FR 44435, Nov. 6, 1984]



Sec. 27.1415  Ditching equipment.

    (a) Emergency flotation and signaling equipment required by any 
operating rule in this chapter must meet the requirements of this 
section.
    (b) Each raft and each life preserver must be approved and must be 
installed so that it is readily available to the crew and passengers. 
The storage provisions for life preservers must accommodate one life 
preserver for each occupant for which certification for ditching is 
requested.
    (c) Each raft released automatically or by the pilot must be 
attached to the rotorcraft by a line to keep it alongside the 
rotorcraft. This line must be weak enough to break before submerging the 
empty raft to which it is attached.
    (d) Each signaling device must be free from hazard in its operation 
and must be installed in an accessible location.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55470, Dec. 20, 1976]



Sec. 27.1419  Ice protection.

    (a) To obtain certification for flight into icing conditions, 
compliance with this section must be shown.
    (b) It must be demonstrated that the rotorcraft can be safely 
operated in the continuous maximum and intermittent maximum icing 
conditions determined under appendix C of Part 29 of this chapter within 
the rotorcraft altitude envelope. An analysis must be performed to 
establish, on the basis of the rotorcraft's operational needs, the 
adequacy of the ice protection system for the various components of the 
rotorcraft.
    (c) In addition to the analysis and physical evaluation prescribed 
in paragraph (b) of this section, the effectiveness of the ice 
protection system and

[[Page 728]]

its components must be shown by flight tests of the rotorcraft or its 
components in measured natural atmospheric icing conditions and by one 
or more of the following tests as found necessary to determine the 
adequacy of the ice protection system:
    (1) Laboratory dry air or simulated icing tests, or a combination of 
both, of the components or models of the components.
    (2) Flight dry air tests of the ice protection system as a whole, or 
its individual components.
    (3) Flight tests of the rotorcraft or its components in measured 
simulated icing conditions.
    (d) The ice protection provisions of this section are considered to 
be applicable primarily to the airframe. Powerplant installation 
requirements are contained in Subpart E of this part.
    (e) A means must be indentified or provided for determining the 
formation of ice on critical parts of the rotorcraft. Unless otherwise 
restricted, the means must be available for nighttime as well as daytime 
operation. The rotorcraft flight manual must describe the means of 
determining ice formation and must contain information necessary for 
safe operation of the rotorcraft in icing conditions.

[Amdt. 27-19, 48 FR 4389, Jan. 31, 1983]



Sec. 27.1435  Hydraulic systems.

    (a) Design. Each hydraulic system and its elements must withstand, 
without yielding, any structural loads expected in addition to hydraulic 
loads.
    (b) Tests. Each system must be substantiated by proof pressure 
tests. When proof tested, no part of any system may fail, malfunction, 
or experience a permanent set. The proof load of each system must be at 
least 1.5 times the maximum operating pressure of that system.
    (c) Accumulators. No hydraulic accumulator or pressurized reservoir 
may be installed on the engine side of any firewall unless it is an 
integral part of an engine.



Sec. 27.1457  Cockpit voice recorders.

    (a) Each cockpit voice recorder required by the operating rules of 
this chapter must be approved, and must be installed so that it will 
record the following:
    (1) Voice communications transmitted from or received in the 
rotorcraft by radio.
    (2) Voice communications of flight crewmembers on the flight deck.
    (3) Voice communications of flight crewmembers on the flight deck, 
using the rotorcraft's interphone system.
    (4) Voice or audio signals identifying navigation or approach aids 
introduced into a headset or speaker.
    (5) Voice communications of flight crewmembers using the passenger 
loudspeaker system, if there is such a system, and if the fourth channel 
is available in accordance with the requirements of paragraph (c)(4)(ii) 
of this section.
    (6) If datalink communication equipment is installed, all datalink 
communications, using an approved data message set. Datalink messages 
must be recorded as the output signal from the communications unit that 
translates the signal into usable data.
    (b) The recording requirements of paragraph (a)(2) of this section 
may be met:
    (1) By installing a cockpit-mounted area microphone located in the 
best position for recording voice communications originating at the 
first and second pilot stations and voice communications of other 
crewmembers on the flight deck when directed to those stations; or
    (2) By installing a continually energized or voice-actuated lip 
microphone at the first and second pilot stations.
    The microphone specified in this paragraph must be so located and, 
if necessary, the preamplifiers and filters of the recorder must be 
adjusted or supplemented so that the recorded communications are 
intelligible when recorded under flight cockpit noise conditions and 
played back. The level of intelligibility must be approved by the 
Administrator. Repeated aural or visual playback of the record may be 
used in evaluating intelligibility.
    (c) Each cockpit voice recorder must be installed so that the part 
of the communication or audio signals specified in paragraph (a) of this 
section obtained from each of the following

[[Page 729]]

sources is recorded on a separate channel:
    (1) For the first channel, from each microphone, headset, or speaker 
used at the first pilot station.
    (2) For the second channel, from each microphone, headset, or 
speaker used at the second pilot station.
    (3) For the third channel, from the cockpit-mounted area microphone, 
or the continually energized or voice-actuated lip microphone at the 
first and second pilot stations.
    (4) For the fourth channel, from:
    (i) Each microphone, headset, or speaker used at the stations for 
the third and fourth crewmembers; or
    (ii) If the stations specified in paragraph (c)(4)(i) of this 
section are not required or if the signal at such a station is picked up 
by another channel, each microphone on the flight deck that is used with 
the passenger loudspeaker system if its signals are not picked up by 
another channel.
    (iii) Each microphone on the flight deck that is used with the 
rotorcraft's loudspeaker system if its signals are not picked up by 
another channel.
    (d) Each cockpit voice recorder must be installed so that:
    (1)(i) It receives its electrical power from the bus that provides 
the maximum reliability for operation of the cockpit voice recorder 
without jeopardizing service to essential or emergency loads.
    (ii) It remains powered for as long as possible without jeopardizing 
emergency operation of the rotorcraft.
    (2) There is an automatic means to simultaneously stop the recorder 
and prevent each erasure feature from functioning, within 10 minutes 
after crash impact;
    (3) There is an aural or visual means for preflight checking of the 
recorder for proper operation;
    (4) Whether the cockpit voice recorder and digital flight data 
recorder are installed in separate boxes or in a combination unit, no 
single electrical failure external to the recorder may disable both the 
cockpit voice recorder and the digital flight data recorder; and
    (5) It has an independent power source--
    (i) That provides 10 1 minutes of electrical 
power to operate both the cockpit voice recorder and cockpit-mounted 
area microphone;
    (ii) That is located as close as practicable to the cockpit voice 
recorder; and
    (iii) To which the cockpit voice recorder and cockpit-mounted area 
microphone are switched automatically in the event that all other power 
to the cockpit voice recorder is interrupted either by normal shutdown 
or by any other loss of power to the electrical power bus.
    (e) The record container must be located and mounted to minimize the 
probability of rupture of the container as a result of crash impact and 
consequent heat damage to the record from fire.
    (f) If the cockpit voice recorder has a bulk erasure device, the 
installation must be designed to minimize the probability of inadvertent 
operation and actuation of the device during crash impact.
    (g) Each recorder container must be either bright orange or bright 
yellow.
    (h) When both a cockpit voice recorder and a flight data recorder 
are required by the operating rules, one combination unit may be 
installed, provided that all other requirements of this section and the 
requirements for flight data recorders under this part are met.

[Amdt. 27-22, 53 FR 26144, July 11, 1988, as amended by Amdt. 27-43, 73 
FR 12563, Mar. 7, 2008; 74 FR 32800, July 9, 2009; Amdt. 27-45, 75 FR 
17045, Apr. 5, 2010]



Sec. 27.1459  Flight data recorders.

    (a) Each flight recorder required by the operating rules of 
Subchapter G of this chapter must be installed so that:
    (1) It is supplied with airspeed, altitude, and directional data 
obtained from sources that meet the accuracy requirements of Sec. Sec. 
27.1323, 27.1325, and 27.1327 of this part, as applicable;
    (2) The vertical acceleration sensor is rigidly attached, and 
located longitudinally within the approved center of gravity limits of 
the rotorcraft;
    (3)(i) It receives its electrical power from the bus that provides 
the maximum reliability for operation of the

[[Page 730]]

flight data recorder without jeopardizing service to essential or 
emergency loads.
    (ii) It remains powered for as long as possible without jeopardizing 
emergency operation of the rotorcraft.
    (4) There is an aural or visual means for preflight checking of the 
recorder for proper recording of data in the storage medium;
    (5) Except for recorders powered solely by the engine-driven 
electrical generator system, there is an automatic means to 
simultaneously stop a recorder that has a data erasure feature and 
prevent each erasure feature from functioning, within 10 minutes after 
any crash impact; and
    (6) Whether the cockpit voice recorder and digital flight data 
recorder are installed in separate boxes or in a combination unit, no 
single electrical failure external to the recorder may disable both the 
cockpit voice recorder and the digital flight data recorder.
    (b) Each nonejectable recorder container must be located and mounted 
so as to minimize the probability of container rupture resulting from 
crash impact and subsequent damage to the record from fire.
    (c) A correlation must be established between the flight recorder 
readings of airspeed, altitude, and heading and the corresponding 
readings (taking into account correction factors) of the first pilot's 
instruments. This correlation must cover the airspeed range over which 
the aircraft is to be operated, the range of altitude to which the 
aircraft is limited, and 360 degrees of heading. Correlation may be 
established on the ground as appropriate.
    (d) Each recorder container must:
    (1) Be either bright orange or bright yellow;
    (2) Have a reflective tape affixed to its external surface to 
facilitate its location under water; and
    (3) Have an underwater locating device, when required by the 
operating rules of this chapter, on or adjacent to the container which 
is secured in such a manner that they are not likely to be separated 
during crash impact.
    (e) When both a cockpit voice recorder and a flight data recorder 
are required by the operating rules, one combination unit may be 
installed, provided that all other requirements of this section and the 
requirements for cockpit voice recorders under this part are met.

[Amdt. 27-22, 53 FR 26144, July 11, 1988, as amended by Amdt. 27-43, 73 
FR 12564, Mar. 7, 2008; 74 FR 32800, July 9, 2009; Amdt. 27-45, 75 FR 
17045, Apr. 5, 2010]



Sec. 27.1461  Equipment containing high energy rotors.

    (a) Equipment containing high energy rotors must meet paragraph (b), 
(c), or (d) of this section.
    (b) High energy rotors contained in equipment must be able to 
withstand damage caused by malfunctions, vibration, abnormal speeds, and 
abnormal temperatures. In addition--
    (1) Auxiliary rotor cases must be able to contain damage caused by 
the failure of high energy rotor blades; and
    (2) Equipment control devices, systems, and instrumentation must 
reasonably ensure that no operating limitations affecting the integrity 
of high energy rotors will be exceeded in service.
    (c) It must be shown by test that equipment containing high energy 
rotors can contain any failure of a high energy rotor that occurs at the 
highest speed obtainable with the normal speed control devices 
inoperative.
    (d) Equipment containing high energy rotors must be located where 
rotor failure will neither endanger the occupants nor adversely affect 
continued safe flight.

[Amdt. 27-2, 33 FR 964, Jan. 26, 1968]



             Subpart G_Operating Limitations and Information



Sec. 27.1501  General.

    (a) Each operating limitation specified in Sec. Sec. 27.1503 
through 27.1525 and other limitations and information necessary for safe 
operation must be established.
    (b) The operating limitations and other information necessary for 
safe operation must be made available to

[[Page 731]]

the crewmembers as prescribed in Sec. Sec. 27.1541 through 27.1589.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]

                          Operating Limitations



Sec. 27.1503  Airspeed limitations: general.

    (a) An operating speed range must be established.
    (b) When airspeed limitations are a function of weight, weight 
distribution, altitude, rotor speed, power, or other factors, airspeed 
limitations corresponding with the critical combinations of these 
factors must be established.



Sec. 27.1505  Never-exceed speed.

    (a) The never-exceed speed, VNE, must be established so that it is--
    (1) Not less than 40 knots (CAS); and
    (2) Not more than the lesser of--
    (i) 0.9 times the maximum forward speeds established under Sec. 
27.309;
    (ii) 0.9 times the maximum speed shown under Sec. Sec. 27.251 and 
27.629; or
    (iii) 0.9 times the maximum speed substantiated for advancing blade 
tip mach number effects.
    (b) VNE may vary with altitude, r.p.m., temperature, and weight, 
if--
    (1) No more than two of these variables (or no more than two 
instruments integrating more than one of these variables) are used at 
one time; and
    (2) The ranges of these variables (or of the indications on 
instruments integrating more than one of these variables) are large 
enough to allow an operationally practical and safe variation of VNE.
    (c) For helicopters, a stabilized power-off VNE denoted as VNE 
(power-off) may be established at a speed less than VNE established 
pursuant to paragraph (a) of this section, if the following conditions 
are met:
    (1) VNE (power-off) is not less than a speed midway between the 
power-on VNE and the speed used in meeting the requirements of--
    (i) Sec. 27.65(b) for single engine helicopters; and
    (ii) Sec. 27.67 for multiengine helicopters.
    (2) VNE (power-off) is--
    (i) A constant airspeed;
    (ii) A constant amount less than power-on VNE; or
    (iii) A constant airspeed for a portion of the altitude range for 
which certification is requested, and a constant amount less than power-
on VNE for the remainder of the altitude range.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 27-2, 33 FR 964, Jan. 26, 1968, and Amdt. 27-14, 43 FR 2325, Jan. 
16, 1978; Amdt. 27-21, 49 FR 44435, Nov. 6, 1984]



Sec. 27.1509  Rotor speed.

    (a) Maximum power-off (autorotation). The maximum power-off rotor 
speed must be established so that it does not exceed 95 percent of the 
lesser of--
    (1) The maximum design r.p.m. determined under Sec. 27.309(b); and
    (2) The maximum r.p.m. shown during the type tests.
    (b) Minimum power off. The minimum power-off rotor speed must be 
established so that it is not less than 105 percent of the greater of--
    (1) The minimum shown during the type tests; and
    (2) The minimum determined by design substantiation.
    (c) Minimum power on. The minimum power-on rotor speed must be 
established so that it is--
    (1) Not less than the greater of--
    (i) The minimum shown during the type tests; and
    (ii) The minimum determined by design substantiation; and
    (2) Not more than a value determined under Sec. 27.33(a)(1) and 
(b)(1).



Sec. 27.1519  Weight and center of gravity.

    The weight and center of gravity limitations determined under 
Sec. Sec. 27.25 and 27.27, respectively, must be established as 
operating limitations.

[Amdt. 27-2, 33 FR 965, Jan. 26, 1968, as amended by Amdt. 27-21, 49 FR 
44435, Nov. 6, 1984]

[[Page 732]]



Sec. 27.1521  Powerplant limitations.

    (a) General. The powerplant limitations prescribed in this section 
must be established so that they do not exceed the corresponding limits 
for which the engines are type certificated.
    (b) Takeoff operation. The powerplant takeoff operation must be 
limited by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value shown during the type tests;
    (2) The maximum allowable manifold pressure (for reciprocating 
engines);
    (3) The time limit for the use of the power corresponding to the 
limitations established in paragraphs (b)(1) and (2) of this section;
    (4) If the time limit in paragraph (b)(3) of this section exceeds 
two minutes, the maximum allowable cylinder head, coolant outlet, or oil 
temperatures;
    (5) The gas temperature limits for turbine engines over the range of 
operating and atmospheric conditions for which certification is 
requested.
    (c) Continuous operation. The continuous operation must be limited 
by--
    (1) The maximum rotational speed which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value shown during the type tests;
    (2) The minimum rotational speed shown under the rotor speed 
requirements in Sec. 27.1509(c); and
    (3) The gas temperature limits for turbine engines over the range of 
operating and atmospheric conditions for which certification is 
requested.
    (d) Fuel grade or designation. The minimum fuel grade (for 
reciprocating engines), or fuel designation (for turbine engines), must 
be established so that it is not less than that required for the 
operation of the engines within the limitations in paragraphs (b) and 
(c) of this section.
    (e) Turboshaft engine torque. For rotorcraft with main rotors driven 
by turboshaft engines, and that do not have a torque limiting device in 
the transmission system, the following apply:
    (1) A limit engine torque must be established if the maximum torque 
that the engine can exert is greater than--
    (i) The torque that the rotor drive system is designed to transmit; 
or
    (ii) The torque that the main rotor assembly is designed to 
withstand in showing compliance with Sec. 27.547(e).
    (2) The limit engine torque established under paragraph (e)(1) of 
this section may not exceed either torque specified in paragraph 
(e)(1)(i) or (ii) of this section.
    (f) Ambient temperature. For turbine engines, ambient temperature 
limitations (including limitations for winterization installations, if 
applicable) must be established as the maximum ambient atmospheric 
temperature at which compliance with the cooling provisions of 
Sec. Sec. 27.1041 through 27.1045 is shown.
    (g) Two and one-half-minute OEI power operation. Unless otherwise 
authorized, the use of 2\1/2\-minute OEI power must be limited to engine 
failure operation of multiengine, turbine-powered rotorcraft for not 
longer than 2\1/2\ minutes after failure of an engine. The use of 2\1/
2\-minute OEI power must also be limited by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum demonstrated during the type tests;
    (2) The maximum allowable gas temperature; and
    (3) The maximum allowable torque.
    (h) Thirty-minute OEI power operation. Unless otherwise authorized, 
the use of 30-minute OEI power must be limited to multiengine, turbine-
powered rotorcraft for not longer than 30 minutes after failure of an 
engine. The use of 30-minute OEI power must also be limited by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value demonstrated during the type tests;
    (2) The maximum allowable gas temperature; and
    (3) The maximum allowable torque.
    (i) Continuous OEI power operation. Unless otherwise authorized, the 
use of

[[Page 733]]

continuous OEI power must be limited to multiengine, turbine-powered 
rotorcraft for continued flight after failure of an engine. The use of 
continuous OEI power must also be limited by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value demonstrated during the type tests;
    (2) The maximum allowable gas temperature; and
    (3) The maximum allowable torque.
    (j) Rated 30-second OEI power operation. Rated 30-second OEI power 
is permitted only on multiengine, turbine-powered rotorcraft, also 
certificated for the use of rated 2-minute OEI power, and can only be 
used for continued operation of the remaining engine(s) after a failure 
or precautionary shutdown of an engine. It must be shown that following 
application of 30-second OEI power, any damage will be readily 
detectable by the applicable inspections and other related procedures 
furnished in accordance with Section A27.4 of appendix A of this part 
and Section A33.4 of appendix A of part 33. The use of 30-second OEI 
power must be limited to not more than 30 seconds for any period in 
which that power is used, and by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value demonstrated during the type tests;
    (2) The maximum allowable gas temperature; and
    (3) The maximum allowable torque.
    (k) Rated 2-minute OEI power operation. Rated 2-minute OEI power is 
permitted only on multiengine, turbine-powered rotorcraft, also 
certificated for the use of rated 30-second OEI power, and can only be 
used for continued operation of the remaining engine(s) after a failure 
or precautionary shutdown of an engine. It must be shown that following 
application of 2-minute OEI power, any damage will be readily detectable 
by the applicable inspections and other related procedures furnished in 
accordance with Section A27.4 of appendix A of this part and Section 
A33.4 of appendix A of part 33. The use of 2-minute OEI power must be 
limited to not more than 2 minutes for any period in which that power is 
used, and by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value demonstrated during the type tests;
    (2) The maximum allowable gas temperature; and
    (3) The maximum allowable torque.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 
43 FR 2325, Jan. 16, 1978; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988; 
Amdt. 27-29, 59 FR 47767, Sept. 16, 1994]



Sec. 27.1523  Minimum flight crew.

    The minimum flight crew must be established so that it is sufficient 
for safe operation, considering--
    (a) The workload on individual crewmembers;
    (b) The accessibility and ease of operation of necessary controls by 
the appropriate crewmember; and
    (c) The kinds of operation authorized under Sec. 27.1525.



Sec. 27.1525  Kinds of operations.

    The kinds of operations (such as VFR, IFR, day, night, or icing) for 
which the rotorcraft is approved are established by demonstrated 
compliance with the applicable certification requirements and by the 
installed equipment.

[Amdt. 27-21, 49 FR 44435, Nov. 6, 1984]



Sec. 27.1527  Maximum operating altitude.

    The maximum altitude up to which operation is allowed, as limited by 
flight, structural, powerplant, functional, or equipment 
characteristics, must be established.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]

[[Page 734]]



Sec. 27.1529  Instructions for Continued Airworthiness.

    The applicant must prepare Instructions for Continued Airworthiness 
in accordance with appendix A to this part that are acceptable to the 
Administrator. The instructions may be incomplete at type certification 
if a program exists to ensure their completion prior to delivery of the 
first rotorcraft or issuance of a standard certificate of airworthiness, 
whichever occurs later.

[Amdt. 27-18, 45 FR 60177, Sept. 11, 1980]

                          Markings and Placards



Sec. 27.1541  General.

    (a) The rotorcraft must contain--
    (1) The markings and placards specified in Sec. Sec. 27.1545 
through 27.1565, and
    (2) Any additional information, instrument markings, and placards 
required for the safe operation of rotorcraft with unusual design, 
operating or handling characteristics.
    (b) Each marking and placard prescribed in paragraph (a) of this 
section--
    (1) Must be displayed in a conspicuous place; and
    (2) May not be easily erased, disfigured, or obscured.



Sec. 27.1543  Instrument markings: general.

    For each instrument--
    (a) When markings are on the cover glass of the instrument, there 
must be means to maintain the correct alignment of the glass cover with 
the face of the dial; and
    (b) Each arc and line must be wide enough, and located, to be 
clearly visible to the pilot.



Sec. 27.1545  Airspeed indicator.

    (a) Each airspeed indicator must be marked as specified in paragraph 
(b) of this section, with the marks located at the corresponding 
indicated airspeeds.
    (b) The following markings must be made:
    (1) A red radial line--
    (i) For rotocraft other than helicopters, at VNE; and
    (ii) For helicopters at VNE (power-on).
    (2) A red cross-hatched radial line at VNE (power-off) for 
helicopters, if VNE (power-off) is less than VNE (power-on).
    (3) For the caution range, a yellow arc.
    (4) For the safe operating range, a green arc.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 
43 FR 2325, Jan. 16, 1978; 43 FR 3900, Jan. 30, 1978; Amdt. 27-16, 43 FR 
50599, Oct. 30, 1978]



Sec. 27.1547  Magnetic direction indicator.

    (a) A placard meeting the requirements of this section must be 
installed on or near the magnetic direction indicator.
    (b) The placard must show the calibration of the instrument in level 
flight with the engines operating.
    (c) The placard must state whether the calibration was made with 
radio receivers on or off.
    (d) Each calibration reading must be in terms of magnetic heading in 
not more than 45 degree increments.
    (e) If a magnetic nonstabilized direction indicator can have a 
deviation of more than 10 degrees caused by the operation of electrical 
equipment, the placard must state which electrical loads, or combination 
of loads, would cause a deviation of more than 10 degrees when turned 
on.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-13, 
42 FR 36972, July 18, 1977]



Sec. 27.1549  Powerplant instruments.

    For each required powerplant instrument, as appropriate to the type 
of instrument--
    (a) Each maximum and, if applicable, minimum safe operating limit 
must be marked with a red radial or a red line;
    (b) Each normal operating range must be marked with a green arc or 
green line, not extending beyond the maximum and minimum safe limits;
    (c) Each takeoff and precautionary range must be marked with a 
yellow arc or yellow line;

[[Page 735]]

    (d) Each engine or propeller range that is restricted because of 
excessive vibration stresses must be marked with red arcs or red lines; 
and
    (e) Each OEI limit or approved operating range must be marked to be 
clearly differentiated from the markings of paragraphs (a) through (d) 
of this section except that no marking is normally required for the 30-
second OEI limit.

[Amdt. 27-11, 41 FR 55470, Dec. 20, 1976, as amended by Amdt. 27-23, 53 
FR 34215, Sept. 2, 1988; Amdt. 27-29, 59 FR 47768, Sept. 16, 1994]



Sec. 27.1551  Oil quantity indicator.

    Each oil quantity indicator must be marked with enough increments to 
indicate readily and accurately the quantity of oil.



Sec. 27.1553  Fuel quantity indicator.

    If the unusable fuel supply for any tank exceeds one gallon, or five 
percent of the tank capacity, whichever is greater, a red arc must be 
marked on its indicator extending from the calibrated zero reading to 
the lowest reading obtainable in level flight.



Sec. 27.1555  Control markings.

    (a) Each cockpit control, other than primary flight controls or 
control whose function is obvious, must be plainly marked as to its 
function and method of operation.
    (b) For powerplant fuel controls--
    (1) Each fuel tank selector control must be marked to indicate the 
position corresponding to each tank and to each existing cross feed 
position;
    (2) If safe operation requires the use of any tanks in a specific 
sequence, that sequence must be marked on, or adjacent to, the selector 
for those tanks; and
    (3) Each valve control for any engine of a multiengine rotorcraft 
must be marked to indicate the position corresponding to each engine 
controlled.
    (c) Usable fuel capacity must be marked as follows:
    (1) For fuel systems having no selector controls, the usable fuel 
capacity of the system must be indicated at the fuel quantity indicator.
    (2) For fuel systems having selector controls, the usable fuel 
capacity available at each selector control position must be indicated 
near the selector control.
    (d) For accessory, auxiliary, and emergency controls--
    (1) Each essential visual position indicator, such as those showing 
rotor pitch or landing gear position, must be marked so that each 
crewmember can determine at any time the position of the unit to which 
it relates; and
    (2) Each emergency control must be red and must be marked as to 
method of operation.
    (e) For rotorcraft incorporating retractable landing gear, the 
maximum landing gear operating speed must be displayed in clear view of 
the pilot.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55470, Dec. 20, 1976; Amdt. 27-21, 49 FR 44435, Nov. 6, 1984]



Sec. 27.1557  Miscellaneous markings and placards.

    (a) Baggage and cargo compartments, and ballast location. Each 
baggage and cargo compartment, and each ballast location must have a 
placard stating any limitations on contents, including weight, that are 
necessary under the loading requirements.
    (b) Seats. If the maximum allowable weight to be carried in a seat 
is less than 170 pounds, a placard stating the lesser weight must be 
permanently attached to the seat structure.
    (c) Fuel and oil filler openings. The following apply:
    (1) Fuel filler openings must be marked at or near the filler cover 
with--
    (i) The word ``fuel'';
    (ii) For reciprocating engine powered rotorcraft, the minimum fuel 
grade;
    (iii) For turbine engine powered rotorcraft, the permissible fuel 
designations; and
    (iv) For pressure fueling systems, the maximum permissible fueling 
supply pressure and the maximum permissible defueling pressure.
    (2) Oil filler openings must be marked at or near the filler cover 
with the word ``oil''.
    (d) Emergency exit placards. Each placard and operating control for 
each emergency exit must be red. A placard

[[Page 736]]

must be near each emergency exit control and must clearly indicate the 
location of that exit and its method of operation.

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 
41 FR 55471, Dec. 20, 1976]



Sec. 27.1559  Limitations placard.

    There must be a placard in clear view of the pilot that specifies 
the kinds of operations (such as VFR, IFR, day, night, or icing) for 
which the rotorcraft is approved.

[Amdt. 27-21, 49 FR 44435, Nov. 6, 1984]



Sec. 27.1561  Safety equipment.

    (a) Each safety equipment control to be operated by the crew in 
emergency, such as controls for automatic liferaft releases, must be 
plainly marked as to its method of operation.
    (b) Each location, such as a locker or compartment, that carries any 
fire extinguishing, signaling, or other life saving equipment, must be 
so marked.



Sec. 27.1565  Tail rotor.

    Each tail rotor must be marked so that its disc is conspicuous under 
normal daylight ground conditions.

[Amdt. 27-2, 33 FR 965, Jan. 26, 1968]

          Rotorcraft Flight Manual and Approved Manual Material



Sec. 27.1581  General.

    (a) Furnishing information. A Rotorcraft Flight Manual must be 
furnished with each rotorcraft, and it must contain the following:
    (1) Information required by Sec. Sec. 27.1583 through 27.1589.
    (2) Other information that is necessary for safe operation because 
of design, operating, or handling characteristics.
    (b) Approved information. Each part of the manual listed in 
Sec. Sec. 27.1583 through 27.1589, that is appropriate to the 
rotorcraft, must be furnished, verified, and approved, and must be 
segregated, identified, and clearly distinguished from each unapproved 
part of that manual.
    (c) [Reserved]
    (d) Table of contents. Each Rotorcraft Flight Manual must include a 
table of contents if the complexity of the manual indicates a need for 
it.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]



Sec. 27.1583  Operating limitations.

    (a) Airspeed and rotor limitations. Information necessary for the 
marking of airspeed and rotor limitations on, or near, their respective 
indicators must be furnished. The significance of each limitation and of 
the color coding must be explained.
    (b) Powerplant limitations. The following information must be 
furnished:
    (1) Limitations required by Sec. 27.1521.
    (2) Explanation of the limitations, when appropriate.
    (3) Information necessary for marking the instruments required by 
Sec. Sec. 27.1549 through 27.1553.
    (c) Weight and loading distribution. The weight and center of 
gravity limits required by Sec. Sec. 27.25 and 27.27, respectively, 
must be furnished. If the variety of possible loading conditions 
warrants, instructions must be included to allow ready observance of the 
limitations.
    (d) Flight crew. When a flight crew of more than one is required, 
the number and functions of the minimum flight crew determined under 
Sec. 27.1523 must be furnished.
    (e) Kinds of operation. Each kind of operation for which the 
rotorcraft and its equipment installations are approved must be listed.
    (f) [Reserved]
    (g) Altitude. The altitude established under Sec. 27.1527 and an 
explanation of the limiting factors must be furnished.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 
FR 965, Jan. 26, 1968; Amdt. 27-14, 43 FR 2325, Jan. 16, 1978; Amdt. 27-
16, 43 FR 50599, Oct. 30, 1978]

[[Page 737]]



Sec. 27.1585  Operating procedures.

    (a) Parts of the manual containing operating procedures must have 
information concerning any normal and emergency procedures and other 
information necessary for safe operation, including takeoff and landing 
procedures and associated airspeeds. The manual must contain any 
pertinent information including--
    (1) The kind of takeoff surface used in the tests and each 
appropriate climbout speed; and
    (2) The kind of landing surface used in the tests and appropriate 
approach and glide airspeeds.
    (b) For multiengine rotorcraft, information identifying each 
operating condition in which the fuel system independence prescribed in 
Sec. 27.953 is necessary for safety must be furnished, together with 
instructions for placing the fuel system in a configuration used to show 
compliance with that section.
    (c) For helicopters for which a VNE (power-off) is established under 
Sec. 27.1505(c), information must be furnished to explain the VNE 
(power-off) and the procedures for reducing airspeed to not more than 
the VNE (power-off) following failure of all engines.
    (d) For each rotorcraft showing compliance with Sec. 27.1353 (g)(2) 
or (g)(3), the operating procedures for disconnecting the battery from 
its charging source must be furnished.
    (e) If the unusable fuel supply in any tank exceeds five percent of 
the tank capacity, or one gallon, whichever is greater, information must 
be furnished which indicates that when the fuel quantity indicator reads 
``zero'' in level flight, any fuel remaining in the fuel tank cannot be 
used safely in flight.
    (f) Information on the total quantity of usable fuel for each fuel 
tank must be furnished.
    (g) The airspeeds and rotor speeds for minimum rate of descent and 
best glide angle as prescribed in Sec. 27.71 must be provided.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 27-1, 32 FR 6914, May 5, 1967, as amended by Amdt. 27-14, 43 FR 
2326, Jan. 16, 1978; Amdt. 27-16, 43 FR 50599, Oct. 30, 1978; Amdt. 27-
21, 49 FR 44435, Nov. 6, 1984]



Sec. 27.1587  Performance information.

    (a) The Rotorcraft Flight Manual must contain the following 
information, determined in accordance with Sec. Sec. 27.49 through 
27.87 and 27.143(c) and (d):
    (1) Enough information to determine the limiting height-speed 
envelope.
    (2) Information relative to--
    (i) The steady rates of climb and descent, in-ground effect and out-
of-ground effect hovering ceilings, together with the corresponding 
airspeeds and other pertinent information including the calculated 
effects of altitude and temperatures;
    (ii) The maximum weight for each altitude and temperature condition 
at which the rotorcraft can safely hover in-ground effect and out-of-
ground effect in winds of not less than 17 knots from all azimuths. 
These data must be clearly referenced to the appropriate hover charts. 
In addition, if there are other combinations of weight, altitude and 
temperature for which performance information is provided and at which 
the rotorcraft cannot land and take off safely with the maximum wind 
value, those portions of the operating envelope and the appropriate safe 
wind conditions must be stated in the Rotorcraft Flight Manual;
    (iii) For reciprocating engine-powered rotorcraft, the maximum 
atmospheric temperature at which compliance with the cooling provisions 
of Sec. Sec. 27.1041 through 27.1045 is shown; and
    (iv) Glide distance as a function of altitude when autorotating at 
the speeds and conditions for minimum rate of descent and best glide as 
determined in Sec. 27.71.
    (b) The Rotorcraft Flight Manual must contain--

[[Page 738]]

    (1) In its performance information section any pertinent information 
concerning the takeoff weights and altitudes used in compliance with 
Sec. 27.51; and
    (2) The horizontal takeoff distance determined in accordance with 
Sec. 27.65(a)(2)(i).

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 
43 FR 2326, Jan. 16, 1978; Amdt. 27-21, 49 FR 44435, Nov. 6, 1984; Amdt. 
27-44, 73 FR 11000, Feb. 29, 2008; 73 FR 33876, June 16, 2008]



Sec. 27.1589  Loading information.

    There must be loading instructions for each possible loading 
condition between the maximum and minimum weights determined under Sec. 
27.25 that can result in a center of gravity beyond any extreme 
prescribed in Sec. 27.27, assuming any probable occupant weights.



  Sec. Appendix A to Part 27--Instructions for Continued Airworthiness

A27.1 General.

    (a) This appendix specifies requirements for the preparation of 
Instructions for Continued Airworthiness as required by Sec. 27.1529.
    (b) The Instructions for Continued Airworthiness for each rotorcraft 
must include the Instructions for Continued Airworthiness for each 
engine and rotor (hereinafter designated `products'), for each appliance 
required by this chapter, and any required information relating to the 
interface of those appliances and products with the rotorcraft. If 
Instructions for Continued Airworthiness are not supplied by the 
manufacturer of an appliance or product installed in the rotorcraft, the 
Instructions for Continued Airworthiness for the rotorcraft must include 
the information essential to the continued airworthiness of the 
rotorcraft.
    (c) The applicant must submit to the FAA a program to show how 
changes to the Instructions for Continued Airworthiness made by the 
applicant or by the manufacturers of products and appliances installed 
in the rotorcraft will be distributed.

A27.2 Format.

    (a) The Instructions for Continued Airworthiness must be in the form 
of a manual or manuals as appropriate for the quantity of data to be 
provided.
    (b) The format of the manual or manuals must provide for a practical 
arrangement.

A27.3 Content.

    The contents of the manual or manuals must be prepared in the 
English language. The Instructions for Continued Airworthiness must 
contain the following manuals or sections, as appropriate, and 
information:
    (a) Rotorcraft maintenance manual or section. (1) Introduction 
information that includes an explanation of the rotorcraft's features 
and data to the extent necessary for maintenance or preventive 
maintenance.
    (2) A description of the rotorcraft and its systems and 
installations including its engines, rotors, and appliances.
    (3) Basic control and operation information describing how the 
rotorcraft components and systems are controlled and how they operate, 
including any special procedures and limitations that apply.
    (4) Servicing information that covers details regarding servicing 
points, capacities of tanks, reservoirs, types of fluids to be used, 
pressures applicable to the various systems, location of access panels 
for inspection and servicing, locations of lubrication points, the 
lubricants to be used, equipment required for servicing, tow 
instructions and limitations, mooring, jacking, and leveling 
information.
    (b) Maintenance instructions. (1) Scheduling information for each 
part of the rotorcraft and its engines, auxiliary power units, rotors, 
accessories, instruments and equipment that provides the recommended 
periods at which they should be cleaned, inspected, adjusted, tested, 
and lubricated, and the degree of inspection, the applicable wear 
tolerances, and work recommended at these periods. However, the 
applicant may refer to an accessory, instrument, or equipment 
manufacturer as the source of this information if the applicant shows 
the item has an exceptionally high degree of complexity requiring 
specialized maintenance techniques, test equipment, or expertise. The 
recommended overhaul periods and necessary cross references to the 
Airworthiness Limitations section of the manual must also be included. 
In addition, the applicant must include an inspection program that 
includes the frequency and extent of the inspections necessary to 
provide for the continued airworthiness of the rotorcraft.
    (2) Troubleshooting information describing problem malfunctions, how 
to recognize those malfunctions, and the remedial action for those 
malfunctions.
    (3) Information describing the order and method of removing and 
replacing products and parts with any necessary precautions to be taken.
    (4) Other general procedural instructions including procedures for 
system testing during ground running, symmetry checks, weighing and 
determining the center of gravity, lifting and shoring, and storage 
limitations.

[[Page 739]]

    (c) Diagrams of structural access plates and information needed to 
gain access for inspections when access plates are not provided.
    (d) Details for the application of special inspection techniques 
including radiographic and ultrasonic testing where such processes are 
specified.
    (e) Information needed to apply protective treatments to the 
structure after inspection.
    (f) All data relative to structural fasteners such as 
identification, discarded recommendations, and torque values.
    (g) A list of special tools needed.

A27.4 Airworthiness Limitations section.

    The Instructions for Continued Airworthiness must contain a section, 
titled Airworthiness Limitations that is segregated and clearly 
distinguishable from the rest of the document. This section must set 
forth each mandatory replacement time, structural inspection interval, 
and related structural inspection procedure required for type 
certification. If the Instructions for Continued Airworthiness consist 
of multiple documents, the section required by this paragraph must be 
included in the principal manual. This section must contain a legible 
statement in a prominent location that reads: ``The Airworthiness 
Limitations section is FAA approved and specifies inspections and other 
maintenance required under Sec. Sec. 43.16 and 91.403 of the Federal 
Aviation Regulations unless an alternative program has been FAA 
approved.''

[Amdt. 27-18, 45 FR 60177, Sept. 11, 1980, as amended by Amdt. 27-24, 54 
FR 34329, Aug. 18, 1989; Amdt. 27-47, 76 FR 74663, Dec. 1, 2011]



   Sec. Appendix B to Part 27--Airworthiness Criteria for Helicopter 
                            Instrument Flight

    I. General. A normal category helicopter may not be type 
certificated for operation under the instrument flight rules (IFR) of 
this chapter unless it meets the design and installation requirements 
contained in this appendix.
    II. Definitions. (a) VYI means instrument climb speed, 
utilized instead of VY for compliance with the climb 
requirements for instrument flight.
    (b) VNEI means instrument flight never exceed speed, 
utilized instead of VNE for compliance with maximum limit 
speed requirements for instrument flight.
    (c) VMINI means instrument flight minimum speed, utilized 
in complying with minimum limit speed requirements for instrument 
flight.
    III. Trim. It must be possible to trim the cyclic, collective, and 
directional control forces to zero at all approved IFR airspeeds, power 
settings, and configurations appropriate to the type.
    IV. Static longitudinal stability. (a) General. The helicopter must 
possess positive static longitudinal control force stability at critical 
combinations of weight and center of gravity at the conditions specified 
in paragraph IV (b) or (c) of this appendix, as appropriate. The stick 
force must vary with speed so that any substantial speed change results 
in a stick force clearly perceptible to the pilot. For single-pilot 
approval, the airspeed must return to within 10 percent of the trim 
speed when the control force is slowly released for each trim condition 
specified in paragraph IV(b) of the this appendix.
    (b) For single-pilot approval:
    (1) Climb. Stability must be shown in climb throughout the speed 
range 20 knots either side of trim with--
    (i) The helicopter trimmed at VYI;
    (ii) Landing gear retracted (if retractable); and
    (iii) Power required for limit climb rate (at least 1,000 fpm) at 
VYI or maximum continuous power, whichever is less.
    (2) Cruise. Stability must be shown throughout the speed range from 
0.7 to 1.1 VH or VNEI, whichever is lower, not to 
exceed 20 knots from trim with--
    (i) The helicopter trimmed and power adjusted for level flight at 
0.9 VH or 0.9 VNEI, whichever is lower; and
    (ii) Landing gear retracted (if retractable).
    (3) Slow cruise. Stability must be shown throughout the speed range 
from 0.9 VMINI to 1.3 VMINI or 20 knots above trim 
speed, whichever is greater, with--
    (i) the helicopter trimmed and power adjusted for level flight at 
1.1 VMINI; and
    (ii) Landing gear retracted (if retractable).
    (4) Descent. Stability must be shown throughout the speed range 20 
knots either side of trim with--
    (i) The helicopter trimmed at 0.8 VH or 0.8 
VNEI (or 0.8 VLE for the landing gear extended 
case), whichever is lower;
    (ii) Power required for 1,000 fpm descent at trim speed; and
    (iii) Landing gear extended and retracted, if applicable.
    (5) Approach. Stability must be shown throughout the speed range 
from 0.7 times the minimum recommended approach speed to 20 knots above 
the maximum recommended approach speed with--
    (i) The helicopter trimmed at the recommended approach speed or 
speeds;
    (ii) Landing gear extended and retracted, if applicable; and
    (iii) Power required to maintain a 3[deg] glide path and power 
required to maintain the steepest approach gradient for which approval 
is requested.
    (c) Helicopters approved for a minimum crew of two pilots must 
comply with the provisions of paragraphs IV(b)(2) and IV(b)(5) of this 
appendix.

[[Page 740]]

    V. Static Lateral Directional Stability. (a) Static directional 
stability must be positive throughout the approved ranges of airspeed, 
power, and vertical speed. In straight and steady sideslips up to 10[deg] from trim, directional control position must 
increase without discontinuity with the angle of sideslip, except for a 
small range of sideslip angles around trim. At greater angles up to the 
maximum sideslip angle appropriate to the type, increased directional 
control position must produce an increased angle of sideslip. It must be 
possible to maintain balanced flight without exceptional pilot skill or 
alertness.
    (b) During sideslips up to 10[deg] from trim 
throughout the approved ranges of airspeed, power, and vertical speed, 
there must be no negative dihedral stability perceptible to the pilot 
through lateral control motion or force. Longitudinal cyclic movement 
with sideslip must not be excessive.
    VI. Dynamic stability. (a) For single-pilot approval--
    (1) Any oscillation having a period of less than 5 seconds must damp 
to \1/2\ amplitude in not more than one cycle.
    (2) Any oscillation having a period of 5 seconds or more but less 
than 10 seconds must damp to \1/2\ amplitude in not more than two 
cycles.
    (3) Any oscillation having a period of 10 seconds or more but less 
than 20 seconds must be damped.
    (4) Any oscillation having a period of 20 seconds or more may not 
achieve double amplitude in less than 20 seconds.
    (5) Any aperiodic response may not achieve double amplitude in less 
than 6 seconds.
    (b) For helicopters approved with a minimum crew of two pilots--
    (1) Any oscillation having a period of less than 5 seconds must damp 
to \1/2\ amplitude in not more than two cycles.
    (2) Any oscillation having a period of 5 seconds or more but less 
than 10 seconds must be damped.
    (3) Any oscillation having a period of 10 seconds or more may not 
achieve double amplitude in less than 10 seconds.
    VII. Stability Augmentation System (SAS).
    (a) If a SAS is used, the reliability of the SAS must be related to 
the effects of its failure. Any SAS failure condition that would prevent 
continued safe flight and landing must be extremely improbable. It must 
be shown that, for any failure condition of the SAS that is not shown to 
be extremely improbable--
    (1) The helicopter is safely controllable when the failure or 
malfunction occurs at any speed or altitude within the approved IFR 
operating limitations; and
    (2) The overall flight characteristics of the helicopter allow for 
prolonged instrument flight without undue pilot effort. Additional 
unrelated probable failures affecting the control system must be 
considered. In addition--
    (i) The controllability and maneuverability requirements in Subpart 
B of this part must be met throughout a practical flight envelope;
    (ii) The flight control, trim, and dynamic stability characteristics 
must not be impaired below a level needed to allow continued safe flight 
and landing; and
    (iii) The static longitudinal and static directional stability 
requirements of Subpart B must be met throughout a practical flight 
envelope.
    (b) The SAS must be designed so that it cannot create a hazardous 
deviation in flight path or produce hazardous loads on the helicopter 
during normal operation or in the event of malfunction or failure, 
assuming corrective action begins within an appropriate period of time. 
Where multiple systems are installed, subsequent malfunction conditions 
must be considered in sequence unless their occurrence is shown to be 
improbable.
    VIII. Equipment, systems, and installation. The basic equipment and 
installation must comply with Sec. Sec. 29.1303, 29.1431, and 29.1433 
through Amendment 29-14, with the following exceptions and additions:
    (a) Flight and Navigation Instruments. (1) A magnetic gyro-stablized 
direction indicator instead of a gyroscopic direction indicator required 
by Sec. 29.1303(h); and
    (2) A standby attitude indicator which meets the requirements of 
Sec. Sec. 29.1303(g)(1) through (7) instead of a rate-of-turn indicator 
required by Sec. 29.1303(g). For two-pilot configurations, one pilot's 
primary indicator may be designated for this purpose. If standby 
batteries are provided, they may be charged from the aircraft electrical 
system if adequate isolation is incorporated.
    (b) Miscellaneous requirements. (1) Instrument systems and other 
systems essential for IFR flight that could be adversely affected by 
icing must be adequately protected when exposed to the continuous and 
intermittent maximum icing conditions defined in appendix C of Part 29 
of this chapter, whether or not the rotorcraft is certificated for 
operation in icing conditions.
    (2) There must be means in the generating system to automatically 
de-energize and disconnect from the main bus any power source developing 
hazardous overvoltage.
    (3) Each required flight instrument using a power supply (electric, 
vacuum, etc.) must have a visual means integral with the instrument to 
indicate the adequacy of the power being supplied.
    (4) When multiple systems performing like functions are required, 
each system must be grouped, routed, and spaced so that physical 
separation between systems is provided to ensure that a single 
malfunction will not adversely affect more than one system.

[[Page 741]]

    (5) For systems that operate the required flight instruments at each 
pilot's station--
    (i) Only the required flight instruments for the first pilot may be 
connected to that operating system;
    (ii) Additional instruments, systems, or equipment may not be 
connected to an operating system for a second pilot unless provisions 
are made to ensure the continued normal functioning of the required 
instruments in the event of any malfunction of the additional 
instruments, systems, or equipment which is not shown to be extremely 
improbable;
    (iii) The equipment, systems, and installations must be designed so 
that one display of the information essential to the safety of flight 
which is provided by the instruments will remain available to a pilot, 
without additional crewmember action, after any single failure or 
combination of failures that is not shown to be extremely improbable; 
and
    (iv) For single-pilot configurations, instruments which require a 
static source must be provided with a means of selecting an alternate 
source and that source must be calibrated.
    IX. Rotorcraft Flight Manual. A Rotorcraft Flight Manual or 
Rotorcraft Flight Manual IFR Supplement must be provided and must 
contain--
    (a) Limitations. The approved IFR flight envelope, the IFR 
flightcrew composition, the revised kinds of operation, and the steepest 
IFR precision approach gradient for which the helicopter is approved;
    (b) Procedures. Required information for proper operation of IFR 
systems and the recommended procedures in the event of stability 
augmentation or electrical system failures; and
    (c) Performance. If VYI differs from VY, climb 
performance at VYI and with maximum continuous power 
throughout the ranges of weight, altitude, and temperature for which 
approval is requested.
    X. Electrical and electronic system lightning protection. For 
regulations concerning lightning protection for electrical and 
electronic systems, see Sec. 27.1316.

[Amdt. 27-19, 48 FR 4389, Jan. 31, 1983, as amended by Amdt. 27-44, 73 
FR 11000, Feb. 29, 2008; Amdt. 27-46, 76 FR 33135, June 8, 2011]



           Sec. Appendix C to Part 27--Criteria for Category A

C27.1 General.

    A small multiengine rotorcraft may not be type certificated for 
Category A operation unless it meets the design installation and 
performance requirements contained in this appendix in addition to the 
requirements of this part.

    C27.2 Applicable part 29 sections. The following sections of part 29 
of this chapter must be met in addition to the requirements of this 
part:

29.45(a) and (b)(2)--General.
29.49(a)--Performance at minimum operating speed.
29.51--Takeoff data: General.
29.53--Takeoff: Category A.
29.55--Takeoff decision point: Category A.
29.59--Takeoff Path: Category A.
29.60--Elevated heliport takeoff path: Category A.
29.61--Takeoff distance: Category A.
29.62--Rejected takeoff: Category A.
29.64--Climb: General.
29.65(a)--Climb: AEO.
29.67(a)--Climb: OEI.
29.75--Landing: General.
29.77--Landing decision point: Category A.
29.79--Landing: Category A.
29.81--Landing distance (Ground level sites): Category A.
29.85--Balked landing: Category A.
29.87(a)--Height-velocity envelope.
29.547(a) and (b)--Main and tail rotor structure.
29.861(a)--Fire protection of structure, controls, and other parts.
29.901(c)--Powerplant: Installation.
29.903(b) (c) and (e)--Engines.
29.908(a)--Cooling fans.
29.917(b) and (c)(1)--Rotor drive system: Design.
29.927(c)(1)--Additional tests.
29.953(a)--Fuel system independence.
29.1027(a)--Transmission and gearboxes: General.
29.1045(a)(1), (b), (c), (d), and (f)--Climb cooling test procedures.
29.1047(a)--Takeoff cooling test procedures.
29.1181(a)--Designated fire zones: Regions included.
29.1187(e)--Drainage and ventilation of fire zones.
29.1189(c)--Shutoff means.
29.1191(a)(1)--Firewalls.
29.1193(e)--Cowling and engine compartment covering.
29.1195(a) and (d)--Fire extinguishing systems (one shot).
29.1197--Fire extinguishing agents.
29.1199--Extinguishing agent containers.
29.1201--Fire extinguishing system materials.
29.1305(a) (6) and (b)--Powerplant instruments.
29.1309(b)(2) (i) and (d)--Equipment, systems, and installations.
29.1323(c)(1)--Airspeed indicating system.
29.1331(b)--Instruments using a power supply.
29.1351(d)(2)--Electrical systems and equipment: General (operation 
          without normal electrical power).
29.1587(a)--Performance information.


[[Page 742]]


    Note: In complying with the paragraphs listed in paragraph C27.2 
above, relevant material in the AC ``Certification of Transport Category 
Rotorcraft'' should be used.

[Doc. No. 28008, 61 FR 21907, May 10, 1996]



 Sec. Appendix D to Part 27--HIRF Environments and Equipment HIRF Test 
                                 Levels

    This appendix specifies the HIRF environments and equipment HIRF 
test levels for electrical and electronic systems under Sec. 27.1317. 
The field strength values for the HIRF environments and laboratory 
equipment HIRF test levels are expressed in root-mean-square units 
measured during the peak of the modulation cycle.
    (a) HIRF environment I is specified in the following table:

                      Table I.--HIRF Environment I
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-2 MHz......................................         50         50
2 MHz-30 MHz......................................        100        100
30 MHz-100 MHz....................................         50         50
100 MHz-400 MHz...................................        100        100
400 MHz-700 MHz...................................        700         50
700 MHz-1 GHz.....................................        700        100
1 GHz-2 GHz.......................................      2,000        200
2 GHz-6 GHz.......................................      3,000        200
6 GHz-8 GHz.......................................      1,000        200
8 GHz-12 GHz......................................      3,000        300
12 GHz-18 GHz.....................................      2,000        200
18 GHz-40 GHz.....................................        600        200
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (b) HIRF environment II is specified in the following table:

                     Table II.--HIRF Environment II
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-500 kHz....................................         20         20
500 kHz-2 MHz.....................................         30         30
2 MHz-30 MHz......................................        100        100
30 MHz-100 MHz....................................         10         10
100 MHz-200 MHz...................................         30         10
200 MHz-400 MHz...................................         10         10
400 MHz-1 GHz.....................................        700         40
1 GHz-2 GHz.......................................      1,300        160
2 GHz-4 GHz.......................................      3,000        120
4 GHz-6 GHz.......................................      3,000        160
6 GHz-8 GHz.......................................        400        170
8 GHz-12 GHz......................................      1,230        230
12 GHz-18 GHz.....................................        730        190
18 GHz-40 GHz.....................................        600        150
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (c) HIRF environment III is specified in the following table:

                    Table III.--HIRF Environment III
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-100 kHz....................................        150        150
100 kHz-400 MHz...................................        200        200
400 MHz-700 MHz...................................        730        200
700 MHz-1 GHz.....................................      1,400        240
1 GHz-2 GHz.......................................      5,000        250
2 GHz-4 GHz.......................................      6,000        490
4 GHz-6 GHz.......................................      7,200        400
6 GHz-8 GHz.......................................      1,100        170
8 GHz-12 GHz......................................      5,000        330
12 GHz-18 GHz.....................................      2,000        330
18 GHz-40 GHz.....................................      1,000        420
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (d) Equipment HIRF Test Level 1.
    (1) From 10 kilohertz (kHz) to 400 megahertz (MHz), use conducted 
susceptibility tests with continuous wave (CW) and 1 kHz square wave 
modulation with 90 percent depth or greater. The conducted 
susceptibility current must start at a minimum of 0.6 milliamperes (mA) 
at 10 kHz, increasing 20 decibels (dB) per frequency decade to a minimum 
of 30 mA at 500 kHz.
    (2) From 500 kHz to 40 MHz, the conducted susceptibility current 
must be at least 30 mA.
    (3) From 40 MHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 30 mA at 40 MHz, decreasing 20 dB per frequency 
decade to a minimum of 3 mA at 400 MHz.
    (4) From 100 MHz to 400 MHz, use radiated susceptibility tests at a 
minimum of 20 volts per meter (V/m) peak with CW and 1 kHz square wave 
modulation with 90 percent depth or greater.
    (5) From 400 MHz to 8 gigahertz (GHz), use radiated susceptibility 
tests at a minimum of 150 V/m peak with pulse modulation of 4 percent 
duty cycle with a 1 kHz pulse repetition frequency. This signal must be 
switched on and off at a rate of 1 Hz with a duty cycle of 50 percent.
    (e) Equipment HIRF Test Level 2. Equipment HIRF test level 2 is HIRF 
environment II in table II of this appendix reduced by acceptable 
aircraft transfer function and attenuation curves. Testing must cover 
the frequency band of 10 kHz to 8 GHz.
    (f) Equipment HIRF Test Level 3. (1) From 10 kHz to 400 MHz, use 
conducted susceptibility tests, starting at a minimum of 0.15 mA at 10 
kHz, increasing 20 dB per frequency decade to a minimum of 7.5 mA at 500 
kHz.
    (2) From 500 kHz to 40 MHz, use conducted susceptibility tests at a 
minimum of 7.5 mA.
    (3) From 40 MHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 7.5 mA at 40 MHz, decreasing 20 dB per 
frequency decade to a minimum of 0.75 mA at 400 MHz.

[[Page 743]]

    (4) From 100 MHz to 8 GHz, use radiated susceptibility tests at a 
minimum of 5 V/m.

[Doc. No. FAA-2006-23657, 72 FR 44027, Aug. 6, 2007]



PART 29_AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT--
Table of Contents



                            Subpart A_General

Sec.
29.1 Applicability.
29.2 Special retroactive requirements.

                            Subpart B_Flight

                                 General

29.21 Proof of compliance.
29.25 Weight limits.
29.27 Center of gravity limits.
29.29 Empty weight and corresponding center of gravity.
29.31 Removable ballast.
29.33 Main rotor speed and pitch limits.

                               Performance

29.45 General.
29.49 Performance at minimum operating speed.
29.51 Takeoff data: general.
29.53 Takeoff: Category A.
29.55 Takeoff decision point (TDP): Category A.
29.59 Takeoff path: Category A.
29.60 Elevated heliport takeoff path: Category A.
29.61 Takeoff distance: Category A.
29.62 Rejected takeoff: Category A.
29.63 Takeoff: Category B.
29.64 Climb: General.
29.65 Climb: All engines operating.
29.67 Climb: One engine inoperative (OEI).
29.71 Helicopter angle of glide: Category B.
29.75 Landing: General.
29.77 Landing Decision Point (LDP): Category A.
29.79 Landing: Category A.
29.81 Landing distance: Category A.
29.83 Landing: Category B.
29.85 Balked landing: Category A.
29.87 Height-velocity envelope.

                         Flight Characteristics

29.141 General.
29.143 Controllability and maneuverability.
29.151 Flight controls.
29.161 Trim control.
29.171 Stability: general.
29.173 Static longitudinal stability.
29.175 Demonstration of static longitudinal stability.
29.177 Static directional stability.
29.181 Dynamic stability: Category A rotorcraft.

                Ground and Water Handling Characteristics

29.231 General.
29.235 Taxiing condition.
29.239 Spray characteristics.
29.241 Ground resonance.

                    Miscellaneous Flight Requirements

29.251 Vibration.

                     Subpart C_Strength Requirements

                                 General

29.301 Loads.
29.303 Factor of safety.
29.305 Strength and deformation.
29.307 Proof of structure.
29.309 Design limitations.

                              Flight Loads

29.321 General.
29.337 Limit maneuvering load factor.
29.339 Resultant limit maneuvering loads.
29.341 Gust loads.
29.351 Yawing conditions.
29.361 Engine torque.

                    Control Surface and System Loads

29.391 General.
29.395 Control system.
29.397 Limit pilot forces and torques.
29.399 Dual control system.
29.411 Ground clearance: tail rotor guard.
29.427 Unsymmetrical loads.

                              Ground Loads

29.471 General.
29.473 Ground loading conditions and assumptions.
29.475 Tires and shock absorbers.
29.477 Landing gear arrangement.
29.479 Level landing conditions.
29.481 Tail-down landing conditions.
29.483 One-wheel landing conditions.
29.485 Lateral drift landing conditions.
29.493 Braked roll conditions.
29.497 Ground loading conditions: landing gear with tail wheels.
29.501 Ground loading conditions: landing gear with skids.
29.505 Ski landing conditions.
29.511 Ground load: unsymmetrical loads on multiple-wheel units.

                               Water Loads

29.519 Hull type rotorcraft: Water-based and amphibian.
29.521 Float landing conditions.

                       Main Component Requirements

29.547 Main and tail rotor structure.
29.549 Fuselage and rotor pylon structures.
29.551 Auxiliary lifting surfaces.

[[Page 744]]

                      Emergency Landing Conditions

29.561 General.
29.562 Emergency landing dynamic conditions.
29.563 Structural ditching provisions.

                           Fatigue Evaluation

29.571 Fatigue tolerance evaluation of metallic structure.
29.573 Damage tolerance and fatigue evaluation of composite rotorcraft 
          structures.

                    Subpart D_Design and Construction

                                 General

29.601 Design.
29.602 Critical parts.
29.603 Materials.
29.605 Fabrication methods.
29.607 Fasteners.
29.609 Protection of structure.
29.610 Lightning and static electricity protection.
29.611 Inspection provisions.
29.613 Material strength properties and design values.
29.619 Special factors.
29.621 Casting factors.
29.623 Bearing factors.
29.625 Fitting factors.
29.629 Flutter and divergence.
29.631 Bird strike.

                                 Rotors

29.653 Pressure venting and drainage of rotor blades.
29.659 Mass balance.
29.661 Rotor blade clearance.
29.663 Ground resonance prevention means.

                             Control Systems

29.671 General.
29.672 Stability augmentation, automatic, and power-operated systems.
29.673 Primary flight controls.
29.674 Interconnected controls.
29.675 Stops.
29.679 Control system locks.
29.681 Limit load static tests.
29.683 Operation tests.
29.685 Control system details.
29.687 Spring devices.
29.691 Autorotation control mechanism.
29.695 Power boost and power-operated control system.

                              Landing Gear

29.723 Shock absorption tests.
29.725 Limit drop test.
29.727 Reserve energy absorption drop test.
29.729 Retracting mechanism.
29.731 Wheels.
29.733 Tires.
29.735 Brakes.
29.737 Skis.

                            Floats and Hulls

29.751 Main float buoyancy.
29.753 Main float design.
29.755 Hull buoyancy.
29.757 Hull and auxiliary float strength.

                   Personnel and Cargo Accommodations

29.771 Pilot compartment.
29.773 Pilot compartment view.
29.775 Windshields and windows.
29.777 Cockpit controls.
29.779 Motion and effect of cockpit controls.
29.783 Doors.
29.785 Seats, berths, litters, safety belts, and harnesses.
29.787 Cargo and baggage compartments.
29.801 Ditching.
29.803 Emergency evacuation.
29.805 Flight crew emergency exits.
29.807 Passenger emergency exits.
29.809 Emergency exit arrangement.
29.811 Emergency exit marking.
29.812 Emergency lighting.
29.813 Emergency exit access.
29.815 Main aisle width.
29.831 Ventilation.
29.833 Heaters.

                             Fire Protection

29.851 Fire extinguishers.
29.853 Compartment interiors.
29.855 Cargo and baggage compartments.
29.859 Combustion heater fire protection.
29.861 Fire protection of structure, controls, and other parts.
29.863 Flammable fluid fire protection.

                             External Loads

29.865 External loads.

                              Miscellaneous

29.871 Leveling marks.
29.873 Ballast provisions.

                          Subpart E_Powerplant

                                 General

29.901 Installation.
29.903 Engines.
29.907 Engine vibration.
29.908 Cooling fans.

                           Rotor Drive System

29.917 Design.
29.921 Rotor brake.
29.923 Rotor drive system and control mechanism tests.
29.927 Additional tests.
29.931 Shafting critical speed.
29.935 Shafting joints.
29.939 Turbine engine operating characteristics.

                               Fuel System

29.951 General.

[[Page 745]]

29.952 Fuel system crash resistance.
29.953 Fuel system independence.
29.954 Fuel system lightning protection.
29.955 Fuel flow.
29.957 Flow between interconnected tanks.
29.959 Unusable fuel supply.
29.961 Fuel system hot weather operation.
29.963 Fuel tanks: general.
29.965 Fuel tank tests.
29.967 Fuel tank installation.
29.969 Fuel tank expansion space.
29.971 Fuel tank sump.
29.973 Fuel tank filler connection.
29.975 Fuel tank vents and carburetor vapor vents.
29.977 Fuel tank outlet.
29.979 Pressure refueling and fueling provisions below fuel level.

                         Fuel System Components

29.991 Fuel pumps.
29.993 Fuel system lines and fittings.
29.995 Fuel valves.
29.997 Fuel strainer or filter.
29.999 Fuel system drains.
29.1001 Fuel jettisoning.

                               Oil System

29.1011 Engines: general.
29.1013 Oil tanks.
29.1015 Oil tank tests.
29.1017 Oil lines and fittings.
29.1019 Oil strainer or filter.
29.1021 Oil system drains.
29.1023 Oil radiators.
29.1025 Oil valves.
29.1027 Transmission and gearboxes: general.

                                 Cooling

29.1041 General.
29.1043 Cooling tests.
29.1045 Climb cooling test procedures.
29.1047 Takeoff cooling test procedures.
29.1049 Hovering cooling test procedures.

                            Induction System

29.1091 Air induction.
29.1093 Induction system icing protection.
29.1101 Carburetor air preheater design.
29.1103 Induction systems ducts and air duct systems.
29.1105 Induction system screens.
29.1107 Inter-coolers and after-coolers.
29.1109 Carburetor air cooling.

                             Exhaust System

29.1121 General.
29.1123 Exhaust piping.
29.1125 Exhaust heat exchangers.

                   Powerplant Controls and Accessories

29.1141 Powerplant controls: general.
29.1142 Auxiliary power unit controls.
29.1143 Engine controls.
29.1145 Ignition switches.
29.1147 Mixture controls.
29.1151 Rotor brake controls.
29.1157 Carburetor air temperature controls.
29.1159 Supercharger controls.
29.1163 Powerplant accessories.
29.1165 Engine ignition systems.

                       Powerplant Fire Protection

29.1181 Designated fire zones: regions included.
29.1183 Lines, fittings, and components.
29.1185 Flammable fluids.
29.1187 Drainage and ventilation of fire zones.
29.1189 Shutoff means.
29.1191 Firewalls.
29.1193 Cowling and engine compartment covering.
29.1194 Other surfaces.
29.1195 Fire extinguishing systems.
29.1197 Fire extinguishing agents.
29.1199 Extinguishing agent containers.
29.1201 Fire extinguishing system materials.
29.1203 Fire detector systems.

                           Subpart F_Equipment

                                 General

29.1301 Function and installation.
29.1303 Flight and navigation instruments.
29.1305 Powerplant instruments.
29.1307 Miscellaneous equipment.
29.1309 Equipment, systems, and installations.
29.1316 Electrical and electronic system lightning protection.
29.1317 High-intensity Radiated Fields (HIRF) Protection.

                        Instruments: Installation

29.1321 Arrangement and visibility.
29.1322 Warning, caution, and advisory lights.
29.1323 Airspeed indicating system.
29.1325 Static pressure and pressure altimeter systems.
29.1327 Magnetic direction indicator.
29.1329 Automatic pilot system.
29.1331 Instruments using a power supply.
29.1333 Instrument systems.
29.1335 Flight director systems.
29.1337 Powerplant instruments.

                    Electrical Systems and Equipment

29.1351 General.
29.1353 Electrical equipment and installations.
29.1355 Distribution system.
29.1357 Circuit protective devices.
29.1359 Electrical system fire and smoke protection.
29.1363 Electrical system tests.

                                 Lights

29.1381 Instrument lights.

[[Page 746]]

29.1383 Landing lights.
29.1385 Position light system installation.
29.1387 Position light system dihedral angles.
29.1389 Position light distribution and intensities.
29.1391 Minimum intensities in the horizontal plane of forward and rear 
          position lights.
29.1393 Minimum intensities in any vertical plane of forward and rear 
          position lights.
29.1395 Maximum intensities in overlapping beams of forward and rear 
          position lights.
29.1397 Color specifications.
29.1399 Riding light.
29.1401 Anticollision light system.

                            Safety Equipment

29.1411 General.
29.1413 Safety belts: passenger warning device.
29.1415 Ditching equipment.
29.1419 Ice protection.

                         Miscellaneous Equipment

29.1431 Electronic equipment.
29.1433 Vacuum systems.
29.1435 Hydraulic systems.
29.1439 Protective breathing equipment.
29.1457 Cockpit voice recorders.
29.1459 Flight data recorders.
29.1461 Equipment containing high energy rotors.

             Subpart G_Operating Limitations and Information

29.1501 General.

                          Operating Limitations

29.1503 Airspeed limitations: general.
29.1505 Never-exceed speed.
29.1509 Rotor speed.
29.1517 Limiting height-speed envelope.
29.1519 Weight and center of gravity.
29.1521 Powerplant limitations.
29.1522 Auxiliary power unit limitations.
29.1523 Minimum flight crew.
29.1525 Kinds of operations.
29.1527 Maximum operating altitude.
29.1529 Instructions for Continued Air worthiness.

                          Markings and Placards

29.1541 General.
29.1543 Instrument markings: general.
29.1545 Airspeed indicator.
29.1547 Magnetic direction indicator.
29.1549 Powerplant instruments.
29.1551 Oil quantity indicator.
29.1553 Fuel quantity indicator.
29.1555 Control markings.
29.1557 Miscellaneous markings and placards.
29.1559 Limitations placard.
29.1561 Safety equipment.
29.1565 Tail rotor.

                        Rotorcraft Flight Manual

29.1581 General.
29.1583 Operating limitations.
29.1585 Operating procedures.
29.1587 Performance information.
29.1589 Loading information.

Appendix A to Part 29--Instructions for Continued Airworthiness
Appendix B to Part 29--Airworthiness Criteria for Helicopter Instrument 
          Flight
Appendix C to Part 29--Icing Certification
Appendix D to Part 29--Criteria for Demonstration of Emergency 
          Evacuation Procedures Under Sec. 29.803
Appendix E to Part 29--HIRF Environments and Equipment HIRF Test Levels

    Authority: 49 U.S.C. 106(g), 40113, 44701-44702, 44704.

    Source: Docket No. 5084, 29 FR 16150, Dec. 3, 1964, unless otherwise 
noted.



                            Subpart A_General



Sec. 29.1  Applicability.

    (a) This part prescribes airworthiness standards for the issue of 
type certificates, and changes to those certificates, for transport 
category rotorcraft.
    (b) Transport category rotorcraft must be certificated in accordance 
with either the Category A or Category B requirements of this part. A 
multiengine rotorcraft may be type certificated as both Category A and 
Category B with appropriate and different operating limitations for each 
category.
    (c) Rotorcraft with a maximum weight greater than 20,000 pounds and 
10 or more passenger seats must be type certificated as Category A 
rotorcraft.
    (d) Rotorcraft with a maximum weight greater than 20,000 pounds and 
nine or less passenger seats may be type certificated as Category B 
rotorcraft provided the Category A requirements of Subparts C, D, E, and 
F of this part are met.
    (e) Rotorcraft with a maximum weight of 20,000 pounds or less but 
with 10 or more passenger seats may be type certificated as Category B 
rotorcraft provided the Category A requirements

[[Page 747]]

of Sec. Sec. 29.67(a)(2), 29.87, 29.1517, and subparts C, D, E, and F 
of this part are met.
    (f) Rotorcraft with a maximum weight of 20,000 pounds or less and 
nine or less passenger seats may be type certificated as Category B 
rotorcraft.
    (g) Each person who applies under Part 21 for a certificate or 
change described in paragraphs (a) through (f) of this section must show 
compliance with the applicable requirements of this part.

[Amdt. 29-21, 48 FR 4391, Jan. 31, 1983, as amended by Amdt. 29-39, 61 
FR 21898, May 10, 1996; 61 FR 33963, July 1, 1996]



Sec. 29.2  Special retroactive requirements.

    For each rotorcraft manufactured after September 16, 1992, each 
applicant must show that each occupant's seat is equipped with a safety 
belt and shoulder harness that meets the requirements of paragraphs (a), 
(b), and (c) of this section.
    (a) Each occupant's seat must have a combined safety belt and 
shoulder harness with a single-point release. Each pilot's combined 
safety belt and shoulder harness must allow each pilot, when seated with 
safety belt and shoulder harness fastened, to perform all functions 
necessary for flight operations. There must be a means to secure belts 
and harnesses, when not in use, to prevent interference with the 
operation of the rotorcraft and with rapid egress in an emergency.
    (b) Each occupant must be protected from serious head injury by a 
safety belt plus a shoulder harness that will prevent the head from 
contacting any injurious object.
    (c) The safety belt and shoulder harness must meet the static and 
dynamic strength requirements, if applicable, specified by the 
rotorcraft type certification basis.
    (d) For purposes of this section, the date of manufacture is 
either--
    (1) The date the inspection acceptance records, or equivalent, 
reflect that the rotorcraft is complete and meets the FAA-Approved Type 
Design Data; or
    (2) The date that the foreign civil airworthiness authority 
certifies the rotorcraft is complete and issues an original standard 
airworthiness certificate, or equivalent, in that country.

[Doc. No. 26078, 56 FR 41052, Aug. 16, 1991]



                            Subpart B_Flight

                                 General



Sec. 29.21  Proof of compliance.

    Each requirement of this subpart must be met at each appropriate 
combination of weight and center of gravity within the range of loading 
conditions for which certification is requested. This must be shown--
    (a) By tests upon a rotorcraft of the type for which certification 
is requested, or by calculations based on, and equal in accuracy to, the 
results of testing; and
    (b) By systematic investigation of each required combination of 
weight and center of gravity, if compliance cannot be reasonably 
inferred from combinations investigated.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44435, Nov. 6, 1984]



Sec. 29.25  Weight limits.

    (a) Maximum weight. The maximum weight (the highest weight at which 
compliance with each applicable requirement of this part is shown) or, 
at the option of the applicant, the highest weight for each altitude and 
for each practicably separable operating condition, such as takeoff, 
enroute operation, and landing, must be established so that it is not 
more than--
    (1) The highest weight selected by the applicant;
    (2) The design maximum weight (the highest weight at which 
compliance with each applicable structural loading condition of this 
part is shown); or
    (3) The highest weight at which compliance with each applicable 
flight requirement of this part is shown.
    (4) For Category B rotorcraft with 9 or less passenger seats, the 
maximum weight, altitude, and temperature at which the rotorcraft can 
safely operate near the ground with the maximum wind velocity determined 
under Sec. 29.143(c) and may include other demonstrated wind velocities 
and azimuths. The operating envelopes must

[[Page 748]]

be stated in the Limitations section of the Rotorcraft Flight Manual.
    (b) Minimum weight. The minimum weight (the lowest weight at which 
compliance with each applicable requirement of this part is shown) must 
be established so that it is not less than--
    (1) The lowest weight selected by the applicant;
    (2) The design minimum weight (the lowest weight at which compliance 
with each structural loading condition of this part is shown); or
    (3) The lowest weight at which compliance with each applicable 
flight requirement of this part is shown.
    (c) Total weight with jettisonable external load. A total weight for 
the rotorcraft with a jettisonable external load attached that is 
greater than the maximum weight established under paragraph (a) of this 
section may be established for any rotorcraft-load combination if--
    (1) The rotorcraft-load combination does not include human external 
cargo,
    (2) Structural component approval for external load operations under 
either Sec. 29.865 or under equivalent operational standards is 
obtained,
    (3) The portion of the total weight that is greater than the maximum 
weight established under paragraph (a) of this section is made up only 
of the weight of all or part of the jettisonable external load,
    (4) Structural components of the rotorcraft are shown to comply with 
the applicable structural requirements of this part under the increased 
loads and stresses caused by the weight increase over that established 
under paragraph (a) of this section, and
    (5) Operation of the rotorcraft at a total weight greater than the 
maximum certificated weight established under paragraph (a) of this 
section is limited by appropriate operating limitations under Sec. 
29.865 (a) and (d) of this part.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55471, Dec. 20, 1976; Amdt. 29-43, 64 FR 43020, Aug. 6, 1999; Amdt. 
29-51, 73 FR 11001, Feb. 29, 2008]



Sec. 29.27  Center of gravity limits.

    The extreme forward and aft centers of gravity and, where critical, 
the extreme lateral centers of gravity must be established for each 
weight established under Sec. 29.25. Such an extreme may not lie 
beyond--
    (a) The extremes selected by the applicant;
    (b) The extremes within which the structure is proven; or
    (c) The extremes within which compliance with the applicable flight 
requirements is shown.

[Amdt. 29-3, 33 FR 965, Jan. 26, 1968]



Sec. 29.29  Empty weight and corresponding center of gravity.

    (a) The empty weight and corresponding center of gravity must be 
determined by weighing the rotorcraft without the crew and payload, but 
with--
    (1) Fixed ballast;
    (2) Unusable fuel; and
    (3) Full operating fluids, including--
    (i) Oil;
    (ii) Hydraulic fluid; and
    (iii) Other fluids required for normal operation of rotorcraft 
systems, except water intended for injection in the engines.
    (b) The condition of the rotorcraft at the time of determining empty 
weight must be one that is well defined and can be easily repeated, 
particularly with respect to the weights of fuel, oil, coolant, and 
installed equipment.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29-15, 43 
FR 2326, Jan. 16, 1978]



Sec. 29.31  Removable ballast.

    Removable ballast may be used in showing compliance with the flight 
requirements of this subpart.



Sec. 29.33  Main rotor speed and pitch limits.

    (a) Main rotor speed limits. A range of main rotor speeds must be 
established that--
    (1) With power on, provides adequate margin to accommodate the 
variations in rotor speed occurring in any appropriate maneuver, and is 
consistent

[[Page 749]]

with the kind of governor or synchronizer used; and
    (2) With power off, allows each appropriate autorotative maneuver to 
be performed throughout the ranges of airspeed and weight for which 
certification is requested.
    (b) Normal main rotor high pitch limit (power on). For rotorcraft, 
except helicopters required to have a main rotor low speed warning under 
paragraph (e) of this section, it must be shown, with power on and 
without exceeding approved engine maximum limitations, that main rotor 
speeds substantially less than the minimum approved main rotor speed 
will not occur under any sustained flight condition. This must be met 
by--
    (1) Appropriate setting of the main rotor high pitch stop;
    (2) Inherent rotorcraft characteristics that make unsafe low main 
rotor speeds unlikely; or
    (3) Adequate means to warn the pilot of unsafe main rotor speeds.
    (c) Normal main rotor low pitch limit (power off). It must be shown, 
with power off, that--
    (1) The normal main rotor low pitch limit provides sufficient rotor 
speed, in any autorotative condition, under the most critical 
combinations of weight and airspeed; and
    (2) It is possible to prevent overspeeding of the rotor without 
exceptional piloting skill.
    (d) Emergency high pitch. If the main rotor high pitch stop is set 
to meet paragraph (b)(1) of this section, and if that stop cannot be 
exceeded inadvertently, additional pitch may be made available for 
emergency use.
    (e) Main rotor low speed warning for helicopters. For each single 
engine helicopter, and each multiengine helicopter that does not have an 
approved device that automatically increases power on the operating 
engines when one engine fails, there must be a main rotor low speed 
warning which meets the following requirements:
    (1) The warning must be furnished to the pilot in all flight 
conditions, including power-on and power-off flight, when the speed of a 
main rotor approaches a value that can jeopardize safe flight.
    (2) The warning may be furnished either through the inherent 
aerodynamic qualities of the helicopter or by a device.
    (3) The warning must be clear and distinct under all conditions, and 
must be clearly distinguishable from all other warnings. A visual device 
that requires the attention of the crew within the cockpit is not 
acceptable by itself.
    (4) If a warning device is used, the device must automatically 
deactivate and reset when the low-speed condition is corrected. If the 
device has an audible warning, it must also be equipped with a means for 
the pilot to manually silence the audible warning before the low-speed 
condition is corrected.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 965, Jan. 26, 1968; Amdt. 29-15, 43 FR 2326, Jan. 16, 1978]

                               Performance



Sec. 29.45  General.

    (a) The performance prescribed in this subpart must be determined--
    (1) With normal piloting skill and;
    (2) Without exceptionally favorable conditions.
    (b) Compliance with the performance requirements of this subpart 
must be shown--
    (1) For still air at sea level with a standard atmosphere and;
    (2) For the approved range of atmospheric variables.
    (c) The available power must correspond to engine power, not 
exceeding the approved power, less--
    (1) Installation losses; and
    (2) The power absorbed by the accessories and services at the values 
for which certification is requested and approved.
    (d) For reciprocating engine-powered rotorcraft, the performance, as 
affected by engine power, must be based on a relative humidity of 80 
percent in a standard atmosphere.
    (e) For turbine engine-powered rotorcraft, the performance, as 
affected by

[[Page 750]]

engine power, must be based on a relative humidity of--
    (1) 80 percent, at and below standard temperature; and
    (2) 34 percent, at and above standard temperature plus 50 [deg]F.

Between these two temperatures, the relative humidity must vary 
linearly.
    (f) For turbine-engine-power rotorcraft, a means must be provided to 
permit the pilot to detemine prior to takeoff that each engine is 
capable of developing the power necessary to achieve the applicable 
rotorcraft performance prescribed in this subpart.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-15, 43 
FR 2326, Jan. 16, 1978; Amdt. 29-24, 49 FR 44436, Nov. 6, 1984]



Sec. 29.49  Performance at minimum operating speed.

    (a) For each Category A helicopter, the hovering performance must be 
determined over the ranges of weight, altitude, and temperature for 
which takeoff data are scheduled--
    (1) With not more than takeoff power;
    (2) With the landing gear extended; and
    (3) At a height consistent with the procedure used in establishing 
the takeoff, climbout, and rejected takeoff paths.
    (b) For each Category B helicopter, the hovering performance must be 
determined over the ranges of weight, altitude, and temperature for 
which certification is requested, with--
    (1) Takeoff power;
    (2) The landing gear extended; and
    (3) The helicopter in ground effect at a height consistent with 
normal takeoff procedures.
    (c) For each helicopter, the out-of-ground effect hovering 
performance must be determined over the ranges of weight, altitude, and 
temperature for which certification is requested with takeoff power.
    (d) For rotorcraft other than helicopters, the steady rate of climb 
at the minimum operating speed must be determined over the ranges of 
weight, altitude, and temperature for which certification is requested 
with--
    (1) Takeoff power; and
    (2) The landing gear extended.

[Doc. No. 24802, 61 FR 21898, May 10, 1996; 61 FR 33963, July 1, 1996]



Sec. 29.51  Takeoff data: general.

    (a) The takeoff data required by Sec. Sec. 29.53, 29.55, 29.59, 
29.60, 29.61, 29.62, 29.63, and 29.67 must be determined--
    (1) At each weight, altitude, and temperature selected by the 
applicant; and
    (2) With the operating engines within approved operating 
limitations.
    (b) Takeoff data must--
    (1) Be determined on a smooth, dry, hard surface; and
    (2) Be corrected to assume a level takeoff surface.
    (c) No takeoff made to determine the data required by this section 
may require exceptional piloting skill or alertness, or exceptionally 
favorable conditions.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-39, 61 
FR 21899, May 10, 1996]



Sec. 29.53  Takeoff: Category A.

    The takeoff performance must be determined and scheduled so that, if 
one engine fails at any time after the start of takeoff, the rotorcraft 
can--
    (a) Return to, and stop safely on, the takeoff area; or
    (b) Continue the takeoff and climbout, and attain a configuration 
and airspeed allowing compliance with Sec. 29.67(a)(2).

[Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 FR 33963, July 1, 1996]



Sec. 29.55  Takeoff decision point (TDP): Category A.

    (a) The TDP is the first point from which a continued takeoff 
capability is assured under Sec. 29.59 and is the last point in the 
takeoff path from which a rejected takeoff is assured within the 
distance determined under Sec. 29.62.
    (b) The TDP must be established in relation to the takeoff path 
using no more than two parameters; e.g., airspeed and height, to 
designate the TDP.

[[Page 751]]

    (c) Determination of the TDP must include the pilot recognition time 
interval following failure of the critical engine.

[Doc. No. 24802, 61 FR 21899, May 10, 1996]



Sec. 29.59  Takeoff path: Category A.

    (a) The takeoff path extends from the point of commencement of the 
takeoff procedure to a point at which the rotorcraft is 1,000 feet above 
the takeoff surface and compliance with Sec. 29.67(a)(2) is shown. In 
addition--
    (1) The takeoff path must remain clear of the height-velocity 
envelope established in accordance with Sec. 29.87;
    (2) The rotorcraft must be flown to the engine failure point; at 
which point, the critical engine must be made inoperative and remain 
inoperative for the rest of the takeoff;
    (3) After the critical engine is made inoperative, the rotorcraft 
must continue to the takeoff decision point, and then attain 
VTOSS;
    (4) Only primary controls may be used while attaining 
VTOSS and while establishing a positive rate of climb. 
Secondary controls that are located on the primary controls may be used 
after a positive rate of climb and VTOSS are established but 
in no case less than 3 seconds after the critical engine is made 
inoperative; and
    (5) After attaining VTOSS and a positive rate of a climb, 
the landing gear may be retracted.
    (b) During the takeoff path determination made in accordance with 
paragraph (a) of this section and after attaining VTOSS and a 
positive rate of climb, the climb must be continued at a speed as close 
as practicable to, but not less than, VTOSS until the 
rotorcraft is 200 feet above the takeoff surface. During this interval, 
the climb performance must meet or exceed that required by Sec. 
29.67(a)(1).
    (c) During the continued takeoff, the rotorcraft shall not descend 
below 15 feet above the takeoff surface when the takeoff decision point 
is above 15 feet.
    (d) From 200 feet above the takeoff surface, the rotorcraft takeoff 
path must be level or positive until a height 1,000 feet above the 
takeoff surface is attained with not less than the rate of climb 
required by Sec. 29.67(a)(2). Any secondary or auxiliary control may be 
used after attaining 200 feet above the takeoff surface.
    (e) Takeoff distance will be determined in accordance with Sec. 
29.61.

[Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 FR 33963, July 1, 1996, 
as amended by Amdt. 29-44, 64 FR 45337, Aug. 19, 1999]



Sec. 29.60  Elevated heliport takeoff path: Category A.

    (a) The elevated heliport takeoff path extends from the point of 
commencement of the takeoff procedure to a point in the takeoff path at 
which the rotorcraft is 1,000 feet above the takeoff surface and 
compliance with Sec. 29.67(a)(2) is shown. In addition--
    (1) The requirements of Sec. 29.59(a) must be met;
    (2) While attaining VTOSS and a positive rate of climb, 
the rotorcraft may descend below the level of the takeoff surface if, in 
so doing and when clearing the elevated heliport edge, every part of the 
rotorcraft clears all obstacles by at least 15 feet;
    (3) The vertical magnitude of any descent below the takeoff surface 
must be determined; and
    (4) After attaining VTOSS and a positive rate of climb, 
the landing gear may be retracted.
    (b) The scheduled takeoff weight must be such that the climb 
requirements of Sec. 29.67 (a)(1) and (a)(2) will be met.
    (c) Takeoff distance will be determined in accordance with Sec. 
29.61.

[Doc. No. 24802, 61 FR 21899, May 10, 1996; 61 FR 33963, July 1, 1996]



Sec. 29.61  Takeoff distance: Category A.

    (a) The normal takeoff distance is the horizontal distance along the 
takeoff path from the start of the takeoff to the point at which the 
rotorcraft attains and remains at least 35 feet above the takeoff 
surface, attains and maintains a speed of at least VTOSS, and 
establishes a positive rate of climb, assuming the critical engine 
failure occurs at the engine failure point prior to the takeoff decision 
point.
    (b) For elevated heliports, the takeoff distance is the horizontal 
distance along the takeoff path from the start of the takeoff to the 
point at which the rotorcraft attains and maintains a speed of at least 
VTOSS and establishes a

[[Page 752]]

positive rate of climb, assuming the critical engine failure occurs at 
the engine failure point prior to the takeoff decision point.

[Doc. No. 24802, 61 FR 21899, May 10, 1996]



Sec. 29.62  Rejected takeoff: Category A.

    The rejected takeoff distance and procedures for each condition 
where takeoff is approved will be established with--
    (a) The takeoff path requirements of Sec. Sec. 29.59 and 29.60 
being used up to the TDP where the critical engine failure is recognized 
and the rotorcraft is landed and brought to a complete stop on the 
takeoff surface;
    (b) The remaining engines operating within approved limits;
    (c) The landing gear remaining extended throughout the entire 
rejected takeoff; and
    (d) The use of only the primary controls until the rotorcraft is on 
the ground. Secondary controls located on the primary control may not be 
used until the rotorcraft is on the ground. Means other than wheel 
brakes may be used to stop the rotorcraft if the means are safe and 
reliable and consistent results can be expected under normal operating 
conditions.

[Doc. No. 24802, 61 FR 21899, May 10, 1996, as amended by Amdt. 29-44, 
64 FR 45337, Aug. 19, 1999]



Sec. 29.63  Takeoff: Category B.

    The horizontal distance required to take off and climb over a 50-
foot obstacle must be established with the most unfavorable center of 
gravity. The takeoff may be begun in any manner if--
    (a) The takeoff surface is defined;
    (b) Adequate safeguards are maintained to ensure proper center of 
gravity and control positions; and
    (c) A landing can be made safely at any point along the flight path 
if an engine fails.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55471, Dec. 20, 1976]



Sec. 29.64  Climb: General.

    Compliance with the requirements of Sec. Sec. 29.65 and 29.67 must 
be shown at each weight, altitude, and temperature within the 
operational limits established for the rotorcraft and with the most 
unfavorable center of gravity for each configuration. Cowl flaps, or 
other means of controlling the engine-cooling air supply, will be in the 
position that provides adequate cooling at the temperatures and 
altitudes for which certification is requested.

[Doc. No. 24802, 61 FR 21900, May 10, 1996]



Sec. 29.65  Climb: All engines operating.

    (a) The steady rate of climb must be determined--
    (1) With maximum continuous power;
    (2) With the landing gear retracted; and
    (3) At Vy for standard sea level conditions and at speeds 
selected by the applicant for other conditions.
    (b) For each Category B rotorcraft except helicopters, the rate of 
climb determined under paragraph (a) of this section must provide a 
steady climb gradient of at least 1:6 under standard sea level 
conditions.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29-15, 43 
FR 2326, Jan. 16, 1978; Amdt. 29-39, 61 FR 21900, May 10, 1996; 61 FR 
33963, July 1, 1996]



Sec. 29.67  Climb: One engine inoperative (OEI).

    (a) For Category A rotorcraft, in the critical takeoff configuration 
existing along the takeoff path, the following apply:
    (1) The steady rate of climb without ground effect, 200 feet above 
the takeoff surface, must be at least 100 feet per minute for each 
weight, altitude, and temperature for which takeoff data are to be 
scheduled with--
    (i) The critical engine inoperative and the remaining engines within 
approved operating limitations, except that for rotorcraft for which the 
use of 30-second/2-minute OEI power is requested, only the 2-minute OEI 
power may be used in showing compliance with this paragraph;
    (ii) The landing gear extended; and
    (iii) The takeoff safety speed selected by the applicant.

[[Page 753]]

    (2) The steady rate of climb without ground effect, 1000 feet above 
the takeoff surface, must be at least 150 feet per minute, for each 
weight, altitude, and temperature for which takeoff data are to be 
scheduled with--
    (i) The critical engine inoperative and the remaining engines at 
maximum continuous power including continuous OEI power, if approved, or 
at 30-minute OEI power for rotorcraft for which certification for use of 
30-minute OEI power is requested;
    (ii) The landing gear retracted; and
    (iii) The speed selected by the applicant.
    (3) The steady rate of climb (or descent) in feet per minute, at 
each altitude and temperature at which the rotorcraft is expected to 
operate and at any weight within the range of weights for which 
certification is requested, must be determined with--
    (i) The critical engine inoperative and the remaining engines at 
maximum continuous power including continuous OEI power, if approved, 
and at 30-minute OEI power for rotorcraft for which certification for 
the use of 30-minute OEI power is requested;
    (ii) The landing gear retracted; and
    (iii) The speed selected by the applicant.
    (b) For multiengine Category B rotorcraft meeting the Category A 
engine isolation requirements, the steady rate of climb (or descent) 
must be determined at the speed for best rate of climb (or minimum rate 
of descent) at each altitude, temperature, and weight at which the 
rotorcraft is expected to operate, with the critical engine inoperative 
and the remaining engines at maximum continuous power including 
continuous OEI power, if approved, and at 30-minute OEI power for 
rotorcraft for which certification for the use of 30-minute OEI power is 
requested.

[Doc. No. 24802, 61 FR 21900, May 10, 1996; 61 FR 33963, July 1, 1996, 
as amended by Amdt. 29-44, 64 FR 45337, Aug. 19, 1999; 64 FR 47563, Aug. 
31, 1999]



Sec. 29.71  Helicopter angle of glide: Category B.

    For each category B helicopter, except multiengine helicopters 
meeting the requirements of Sec. 29.67(b) and the powerplant 
installation requirements of category A, the steady angle of glide must 
be determined in autorotation--
    (a) At the forward speed for minimum rate of descent as selected by 
the applicant;
    (b) At the forward speed for best glide angle;
    (c) At maximum weight; and
    (d) At the rotor speed or speeds selected by the applicant.

[Amdt. 29-12, 41 FR 55471, Dec. 20, 1976]



Sec. 29.75  Landing: General.

    (a) For each rotorcraft--
    (1) The corrected landing data must be determined for a smooth, dry, 
hard, and level surface;
    (2) The approach and landing must not require exceptional piloting 
skill or exceptionally favorable conditions; and
    (3) The landing must be made without excessive vertical acceleration 
or tendency to bounce, nose over, ground loop, porpoise, or water loop.
    (b) The landing data required by Sec. Sec. 29.77, 29.79, 29.81, 
29.83, and 29.85 must be determined--
    (1) At each weight, altitude, and temperature for which landing data 
are approved;
    (2) With each operating engine within approved operating 
limitations; and
    (3) With the most unfavorable center of gravity.

[Doc. No. 24802, 61 FR 21900, May 10, 1996]



Sec. 29.77  Landing Decision Point (LDP): Category A.

    (a) The LDP is the last point in the approach and landing path from 
which a balked landing can be accomplished in accordance with Sec. 
29.85.
    (b) Determination of the LDP must include the pilot recognition time 
interval following failure of the critical engine.

[Doc. No. 24802, 64 FR 45338, Aug. 19, 1999]



Sec. 29.79  Landing: Category A.

    (a) For Category A rotorcraft--
    (1) The landing performance must be determined and scheduled so that 
if the critical engine fails at any point in the approach path, the 
rotorcraft can either land and stop safely or climb out and attain a 
rotorcraft configuration

[[Page 754]]

and speed allowing compliance with the climb requirement of Sec. 
29.67(a)(2);
    (2) The approach and landing paths must be established with the 
critical engine inoperative so that the transition between each stage 
can be made smoothly and safely;
    (3) The approach and landing speeds must be selected by the 
applicant and must be appropriate to the type of rotorcraft; and
    (4) The approach and landing path must be established to avoid the 
critical areas of the height-velocity envelope determined in accordance 
with Sec. 29.87.
    (b) It must be possible to make a safe landing on a prepared landing 
surface after complete power failure occurring during normal cruise.

[Doc. No. 24802, 61 FR 21900, May 10, 1996]



Sec. 29.81  Landing distance: Category A.

    The horizontal distance required to land and come to a complete stop 
(or to a speed of approximately 3 knots for water landings) from a point 
50 ft above the landing surface must be determined from the approach and 
landing paths established in accordance with Sec. 29.79.

[Doc. No. 24802, 64 FR 45338, Aug. 19, 1999]



Sec. 29.83  Landing: Category B.

    (a) For each Category B rotorcraft, the horizontal distance required 
to land and come to a complete stop (or to a speed of approximately 3 
knots for water landings) from a point 50 feet above the landing surface 
must be determined with--
    (1) Speeds appropriate to the type of rotorcraft and chosen by the 
applicant to avoid the critical areas of the height-velocity envelope 
established under Sec. 29.87; and
    (2) The approach and landing made with power on and within approved 
limits.
    (b) Each multiengined Category B rotorcraft that meets the 
powerplant installation requirements for Category A must meet the 
requirements of--
    (1) Sections 29.79 and 29.81; or
    (2) Paragraph (a) of this section.
    (c) It must be possible to make a safe landing on a prepared landing 
surface if complete power failure occurs during normal cruise.

[Doc. No. 24802, 61 FR 21900, May 10, 1996; 61 FR 33963, July 1, 1996]



Sec. 29.85  Balked landing: Category A.

    For Category A rotorcraft, the balked landing path with the critical 
engine inoperative must be established so that--
    (a) The transition from each stage of the maneuver to the next stage 
can be made smoothly and safely;
    (b) From the LDP on the approach path selected by the applicant, a 
safe climbout can be made at speeds allowing compliance with the climb 
requirements of Sec. 29.67(a)(1) and (2); and
    (c) The rotorcraft does not descend below 15 feet above the landing 
surface. For elevated heliport operations, descent may be below the 
level of the landing surface provided the deck edge clearance of Sec. 
29.60 is maintained and the descent (loss of height) below the landing 
surface is determined.

[Doc. No. 24802, 64 FR 45338, Aug. 19, 1999]



Sec. 29.87  Height-velocity envelope.

    (a) If there is any combination of height and forward velocity 
(including hover) under which a safe landing cannot be made after 
failure of the critical engine and with the remaining engines (where 
applicable) operating within approved limits, a height-velocity envelope 
must be established for--
    (1) All combinations of pressure altitude and ambient temperature 
for which takeoff and landing are approved; and
    (2) Weight from the maximum weight (at sea level) to the highest 
weight approved for takeoff and landing at each altitude. For 
helicopters, this weight need not exceed the highest weight allowing 
hovering out-of-ground effect at each altitude.
    (b) For single-engine or multiengine rotorcraft that do not meet the 
Category A engine isolation requirements, the height-velocity envelope 
for complete power failure must be established.

[Doc. No. 24802, 61 FR 21901, May 10, 1996; 61 FR 33963, July 1, 1996]

[[Page 755]]

                         Flight Characteristics



Sec. 29.141  General.

    The rotorcraft must--
    (a) Except as specifically required in the applicable section, meet 
the flight characteristics requirements of this subpart--
    (1) At the approved operating altitudes and temperatures;
    (2) Under any critical loading condition within the range of weights 
and centers of gravity for which certification is requested; and
    (3) For power-on operations, under any condition of speed, power, 
and rotor r.p.m. for which certification is requested; and
    (4) For power-off operations, under any condition of speed, and 
rotor r.p.m. for which certification is requested that is attainable 
with the controls rigged in accordance with the approved rigging 
instructions and tolerances;
    (b) Be able to maintain any required flight condition and make a 
smooth transition from any flight condition to any other flight 
condition without exceptional piloting skill, alertness, or strength, 
and without danger of exceeding the limit load factor under any 
operating condition probable for the type, including--
    (1) Sudden failure of one engine, for multiengine rotorcraft meeting 
Transport Category A engine isolation requirements;
    (2) Sudden, complete power failure, for other rotorcraft; and
    (3) Sudden, complete control system failures specified in Sec. 
29.695 of this part; and
    (c) Have any additional characteristics required for night or 
instrument operation, if certification for those kinds of operation is 
requested. Requirements for helicopter instrument flight are contained 
in appendix B of this part.

[Doc. No. 5084, 29 FR 16150, Dec. 8, 1964, as amended by Amdt. 29-3, 33 
FR 905, Jan. 26, 1968; Amdt. 29-12, 41 FR 55471, Dec. 20, 1976; Amdt. 
29-21, 48 FR 4391, Jan. 31, 1983; Amdt. 29-24, 49 FR 44436, Nov. 6, 
1984]



Sec. 29.143  Controllability and maneuverability.

    (a) The rotorcraft must be safely controllable and maneuverable--
    (1) During steady flight; and
    (2) During any maneuver appropriate to the type, including--
    (i) Takeoff;
    (ii) Climb;
    (iii) Level flight;
    (iv) Turning flight;
    (v) Autorotation; and
    (vi) Landing (power on and power off).
    (b) The margin of cyclic control must allow satisfactory roll and 
pitch control at VNE with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Critical rotor r.p.m.; and
    (4) Power off (except for helicopters demonstrating compliance with 
paragraph (f) of this section) and power on.
    (c) Wind velocities from zero to at least 17 knots, from all 
azimuths, must be established in which the rotorcraft can be operated 
without loss of control on or near the ground in any maneuver 
appropriate to the type (such as crosswind takeoffs, sideward flight, 
and rearward flight), with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Critical rotor r.p.m.; and
    (4) Altitude, from standard sea level conditions to the maximum 
takeoff and landing altitude capability of the rotorcraft.
    (d) Wind velocities from zero to at least 17 knots, from all 
azimuths, must be established in which the rotorcraft can be operated 
without loss of control out-of-ground effect, with--
    (1) Weight selected by the applicant;
    (2) Critical center of gravity;
    (3) Rotor r.p.m. selected by the applicant; and
    (4) Altitude, from standard sea level conditions to the maximum 
takeoff and landing altitude capability of the rotorcraft.
    (e) The rotorcraft, after (1) failure of one engine, in the case of 
multiengine rotorcraft that meet Transport Category A engine isolation 
requirements, or (2) complete power failure in the case of other 
rotorcraft, must be controllable over the range of speeds and altitudes 
for which certification is requested when such power failure occurs with 
maximum continuous power and critical weight. No corrective action

[[Page 756]]

time delay for any condition following power failure may be less than--
    (i) For the cruise condition, one second, or normal pilot reaction 
time (whichever is greater); and
    (ii) For any other condition, normal pilot reaction time.
    (f) For helicopters for which a VNE (power-off) is established under 
Sec. 29.1505(c), compliance must be demonstrated with the following 
requirements with critical weight, critical center of gravity, and 
critical rotor r.p.m.:
    (1) The helicopter must be safely slowed to VNE (power-off), without 
exceptional pilot skill after the last operating engine is made 
inoperative at power-on VNE.
    (2) At a speed of 1.1 VNE (power-off), the margin of cyclic control 
must allow satisfactory roll and pitch control with power off.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 965, Jan. 26, 1968; Amdt. 29-15, 43 FR 2326, Jan. 16, 1978; Amdt. 29-
24, 49 FR 44436, Nov. 6, 1984; Amdt. 29-51, 73 FR 11001, Feb. 29, 2008]



Sec. 29.151  Flight controls.

    (a) Longitudinal, lateral, directional, and collective controls may 
not exhibit excessive breakout force, friction, or preload.
    (b) Control system forces and free play may not inhibit a smooth, 
direct rotorcraft response to control system input.

[Amdt. 29-24, 49 FR 44436, Nov. 6, 1984]



Sec. 29.161  Trim control.

    The trim control--
    (a) Must trim any steady longitudinal, lateral, and collective 
control forces to zero in level flight at any appropriate speed; and
    (b) May not introduce any undesirable discontinuities in control 
force gradients.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44436, Nov. 6, 1984]



Sec. 29.171  Stability: general.

    The rotorcraft must be able to be flown, without undue pilot fatigue 
or strain, in any normal maneuver for a period of time as long as that 
expected in normal operation. At least three landings and takeoffs must 
be made during this demonstration.



Sec. 29.173  Static longitudinal stability.

    (a) The longitudinal control must be designed so that a rearward 
movement of the control is necessary to obtain an airspeed less than the 
trim speed, and a forward movement of the control is necessary to obtain 
an airspeed more than the trim speed.
    (b) Throughout the full range of altitude for which certification is 
requested, with the throttle and collective pitch held constant during 
the maneuvers specified in Sec. 29.175(a) through (d), the slope of the 
control position versus airspeed curve must be positive. However, in 
limited flight conditions or modes of operation determined by the 
Administrator to be acceptable, the slope of the control position versus 
airspeed curve may be neutral or negative if the rotorcraft possesses 
flight characteristics that allow the pilot to maintain airspeed within 
5 knots of the desired trim airspeed without 
exceptional piloting skill or alertness.

[Amdt. 29-24, 49 FR 44436, Nov. 6, 1984, as amended by Amdt. 29-51, 73 
FR 11001, Feb. 29, 2008]



Sec. 29.175  Demonstration of static longitudinal stability.

    (a) Climb. Static longitudinal stability must be shown in the climb 
condition at speeds from Vy - 10 kt to Vy + 10 kt with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Maximum continuous power;
    (4) The landing gear retracted; and
    (5) The rotorcraft trimmed at Vy.
    (b) Cruise. Static longitudinal stability must be shown in the 
cruise condition at speeds from 0.8 VNE-10 kt to 0.8 
VNE + 10 kt or, if VH is less than 0.8 
VNE, from VH - 10 kt to VH + 10 kt, with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Power for level flight at 0.8 VNE or VH, 
whichever is less;

[[Page 757]]

    (4) The landing gear retracted; and
    (5) The rotorcraft trimmed at 0.8 VNE or VH, 
whichever is less.
    (c) VNE. Static longitudinal stability must be shown at speeds from 
VNE - 20 kt to VNE with--
    (1) Critical weight;
    (2) Critical center of gravity;
    (3) Power required for level flight at VNE - 10 kt or 
maximum continuous power, whichever is less;
    (4) The landing gear retracted; and
    (5) The rotorcraft trimmed at VNE - 10 kt.
    (d) Autorotation. Static longitudinal stability must be shown in 
autorotation at--
    (1) Airspeeds from the minimum rate of descent airspeed - 10 kt to 
the minimum rate of descent airspeed + 10 kt, with--
    (i) Critical weight;
    (ii) Critical center of gravity;
    (iii) The landing gear extended; and
    (iv) The rotorcraft trimmed at the minimum rate of descent airspeed.
    (2) Airspeeds from the best angle-of-glide airspeed - 10kt to the 
best angle-of-glide airspeed + 10kt, with--
    (i) Critical weight;
    (ii) Critical center of gravity;
    (iii) The landing gear retracted; and
    (iv) The rotorcraft trimmed at the best angle-of-glide airspeed.

[Amdt. 29-51, 73 FR 11001, Feb. 29, 2008]



Sec. 29.177  Static directional stability.

    (a) The directional controls must operate in such a manner that the 
sense and direction of motion of the rotorcraft following control 
displacement are in the direction of the pedal motion with throttle and 
collective controls held constant at the trim conditions specified in 
Sec. 29.175(a), (b), (c), and (d). Sideslip angles must increase with 
steadily increasing directional control deflection for sideslip angles 
up to the lesser of--
    (1) 25 degrees from trim at a speed of 15 
knots less than the speed for minimum rate of descent varying linearly 
to 10 degrees from trim at VNE;
    (2) The steady-state sideslip angles established by Sec. 29.351;
    (3) A sideslip angle selected by the applicant, which corresponds to 
a sideforce of at least 0.1g; or
    (4) The sideslip angle attained by maximum directional control 
input.
    (b) Sufficient cues must accompany the sideslip to alert the pilot 
when approaching sideslip limits.
    (c) During the maneuver specified in paragraph (a) of this section, 
the sideslip angle versus directional control position curve may have a 
negative slope within a small range of angles around trim, provided the 
desired heading can be maintained without exceptional piloting skill or 
alertness.

[Amdt. 29-51, 73 FR 11001, Feb. 29, 2008]



Sec. 29.181  Dynamic stability: Category A rotorcraft.

    Any short-period oscillation occurring at any speed from 
VY to VNE must be positively damped with the 
primary flight controls free and in a fixed position.

[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]

                Ground and Water Handling Characteristics



Sec. 29.231  General.

    The rotorcraft must have satisfactory ground and water handling 
characteristics, including freedom from uncontrollable tendencies in any 
condition expected in operation.



Sec. 29.235  Taxiing condition.

    The rotorcraft must be designed to withstand the loads that would 
occur when the rotorcraft is taxied over the roughest ground that may 
reasonably be expected in normal operation.



Sec. 29.239  Spray characteristics.

    If certification for water operation is requested, no spray 
characteristics during taxiing, takeoff, or landing may obscure the 
vision of the pilot or damage the rotors, propellers, or other parts of 
the rotorcraft.



Sec. 29.241  Ground resonance.

    The rotorcraft may have no dangerous tendency to oscillate on the 
ground with the rotor turning.

[[Page 758]]

                    Miscellaneous Flight Requirements



Sec. 29.251  Vibration.

    Each part of the rotorcraft must be free from excessive vibration 
under each appropriate speed and power condition.



                     Subpart C_Strength Requirements

                                 General



Sec. 29.301  Loads.

    (a) Strength requirements are specified in terms of limit loads (the 
maximum loads to be expected in service) and ultimate loads (limit loads 
multiplied by prescribed factors of safety). Unless otherwise provided, 
prescribed loads are limit loads.
    (b) Unless otherwise provided, the specified air, ground, and water 
loads must be placed in equilibrium with inertia forces, considering 
each item of mass in the rotorcraft. These loads must be distributed to 
closely approximate or conservatively represent actual conditions.
    (c) If deflections under load would significantly change the 
distribution of external or internal loads, this redistribution must be 
taken into account.



Sec. 29.303  Factor of safety.

    Unless otherwise provided, a factor of safety of 1.5 must be used. 
This factor applies to external and inertia loads unless its application 
to the resulting internal stresses is more conservative.



Sec. 29.305  Strength and deformation.

    (a) The structure must be able to support limit loads without 
detrimental or permanent deformation. At any load up to limit loads, the 
deformation may not interfere with safe operation.
    (b) The structure must be able to support ultimate loads without 
failure. This must be shown by--
    (1) Applying ultimate loads to the structure in a static test for at 
least three seconds; or
    (2) Dynamic tests simulating actual load application.



Sec. 29.307  Proof of structure.

    (a) Compliance with the strength and deformation requirements of 
this subpart must be shown for each critical loading condition 
accounting for the environment to which the structure will be exposed in 
operation. Structural analysis (static or fatigue) may be used only if 
the structure conforms to those structures for which experience has 
shown this method to be reliable. In other cases, substantiating load 
tests must be made.
    (b) Proof of compliance with the strength requirements of this 
subpart must include--
    (1) Dynamic and endurance tests of rotors, rotor drives, and rotor 
controls;
    (2) Limit load tests of the control system, including control 
surfaces;
    (3) Operation tests of the control system;
    (4) Flight stress measurement tests;
    (5) Landing gear drop tests; and
    (6) Any additional tests required for new or unusual design 
features.

(Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, 1425)

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-4, 33 
FR 14106, Sept. 18, 1968; Amdt. 27-26, 55 FR 8001, Mar. 6, 1990]



Sec. 29.309  Design limitations.

    The following values and limitations must be established to show 
compliance with the structural requirements of this subpart:
    (a) The design maximum and design minimum weights.
    (b) The main rotor r.p.m. ranges, power on and power off.
    (c) The maximum forward speeds for each main rotor r.p.m. within the 
ranges determined under paragraph (b) of this section.
    (d) The maximum rearward and sideward flight speeds.
    (e) The center of gravity limits corresponding to the limitations 
determined under paragraphs (b), (c), and (d) of this section.
    (f) The rotational speed ratios between each powerplant and each 
connected rotating component.
    (g) The positive and negative limit maneuvering load factors.

[[Page 759]]

                              Flight Loads



Sec. 29.321  General.

    (a) The flight load factor must be assumed to act normal to the 
longitudinal axis of the rotorcraft, and to be equal in magnitude and 
opposite in direction to the rotorcraft inertia load factor at the 
center of gravity.
    (b) Compliance with the flight load requirements of this subpart 
must be shown--
    (1) At each weight from the design minimum weight to the design 
maximum weight; and
    (2) With any practical distribution of disposable load within the 
operating limitations in the Rotorcraft Flight Manual.



Sec. 29.337  Limit maneuvering load factor.

    The rotorcraft must be designed for--
    (a) A limit maneuvering load factor ranging from a positive limit of 
3.5 to a negative limit of -1.0; or
    (b) Any positive limit maneuvering load factor not less than 2.0 and 
any negative limit maneuvering load factor of not less than -0.5 for 
which--
    (1) The probability of being exceeded is shown by analysis and 
flight tests to be extremely remote; and
    (2) The selected values are appropriate to each weight condition 
between the design maximum and design minimum weights.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 27-26, 55 
FR 8002, Mar. 6, 1990]



Sec. 29.339  Resultant limit maneuvering loads.

    The loads resulting from the application of limit maneuvering load 
factors are assumed to act at the center of each rotor hub and at each 
auxiliary lifting surface, and to act in directions and with 
distributions of load among the rotors and auxiliary lifting surfaces, 
so as to represent each critical maneuvering condition, including power-
on and power-off flight with the maximum design rotor tip speed ratio. 
The rotor tip speed ratio is the ratio of the rotorcraft flight velocity 
component in the plane of the rotor disc to the rotational tip speed of 
the rotor blades, and is expressed as follows:
[GRAPHIC] [TIFF OMITTED] TC28SE91.087

where--

V=The airspeed along the flight path (f.p.s.);
a=The angle between the projection, in the plane of symmetry, of the 
          axis of no feathering and a line perpendicular to the flight 
          path (radians, positive when axis is pointing aft);
[Omega]=The angular velocity of rotor (radians per second); and
R=The rotor radius (ft.).



Sec. 29.341  Gust loads.

    Each rotorcraft must be designed to withstand, at each critical 
airspeed including hovering, the loads resulting from vertical and 
horizontal gusts of 30 feet per second.



Sec. 29.351  Yawing conditions.

    (a) Each rotorcraft must be designed for the loads resulting from 
the maneuvers specified in paragraphs (b) and (c) of this section, 
with--
    (1) Unbalanced aerodynamic moments about the center of gravity which 
the aircraft reacts to in a rational or conservative manner considering 
the principal masses furnishing the reacting inertia forces; and
    (2) Maximum main rotor speed.
    (b) To produce the load required in paragraph (a) of this section, 
in unaccelerated flight with zero yaw, at forward speeds from zero up to 
0.6 VNE--
    (1) Displace the cockpit directional control suddenly to the maximum 
deflection limited by the control stops or by the maximum pilot force 
specified in Sec. 29.397(a);
    (2) Attain a resulting sideslip angle or 90[deg], whichever is less; 
and
    (3) Return the directional control suddenly to neutral.
    (c) To produce the load required in paragraph (a) of the section, in 
unaccelerated flight with zero yaw, at forward speeds from 0.6 
VNE up to VNE or VH, whichever is 
less--
    (1) Displace the cockpit directional control suddenly to the maximum 
deflection limited by the control stops or by the maximum pilot force 
specified in Sec. 29.397(a);
    (2) Attain a resulting sideslip angle or 15[deg], whichever is less, 
at the lesser speed of VNE or VH;

[[Page 760]]

    (3) Vary the sideslip angles of paragraphs (b)(2) and (c)(2) of this 
section directly with speed; and
    (4) Return the directional control suddenly to neutral.

[Amdt. 29-26, 55 FR 8002, Mar. 6, 1990, as amended by Amdt. 29-41, 62 FR 
46173, Aug. 29, 1997]



Sec. 29.361  Engine torque.

    The limit engine torque may not be less than the following:
    (a) For turbine engines, the highest of--
    (1) The mean torque for maximum continuous power multiplied by 1.25;
    (2) The torque required by Sec. 29.923;
    (3) The torque required by Sec. 29.927; or
    (4) The torque imposed by sudden engine stoppage due to malfunction 
or structural failure (such as compressor jamming).
    (b) For reciprocating engines, the mean torque for maximum 
continuous power multiplied by--
    (1) 1.33, for engines with five or more cylinders; and
    (2) Two, three, and four, for engines with four, three, and two 
cylinders, respectively.

[Amdt. 29-26, 53 FR 34215, Sept. 2, 1988]

                    Control Surface and System Loads



Sec. 29.391  General.

    Each auxiliary rotor, each fixed or movable stabilizing or control 
surface, and each system operating any flight control must meet the 
requirements of Sec. Sec. 29.395 through 29.399, 29.411, and 29.427.

[Amdt. 29-26, 55 FR 8002, Mar. 6, 1990, as amended by Amdt. 29-41, 62 FR 
46173, Aug. 29, 1997]



Sec. 29.395  Control system.

    (a) The reaction to the loads prescribed in Sec. 29.397 must be 
provided by--
    (1) The control stops only;
    (2) The control locks only;
    (3) The irreversible mechanism only (with the mechanism locked and 
with the control surface in the critical positions for the effective 
parts of the system within its limit of motion);
    (4) The attachment of the control system to the rotor blade pitch 
control horn only (with the control in the critical positions for the 
affected parts of the system within the limits of its motion); and
    (5) The attachment of the control system to the control surface horn 
(with the control in the critical positions for the affected parts of 
the system within the limits of its motion).
    (b) Each primary control system, including its supporting structure, 
must be designed as follows:
    (1) The system must withstand loads resulting from the limit pilot 
forces prescribed in Sec. 29.397;
    (2) Notwithstanding paragraph (b)(3) of this section, when power-
operated actuator controls or power boost controls are used, the system 
must also withstand the loads resulting from the limit pilot forces 
prescribed in Sec. 29.397 in conjunction with the forces output of each 
normally energized power device, including any single power boost or 
actuator system failure;
    (3) If the system design or the normal operating loads are such that 
a part of the system cannot react to the limit pilot forces prescribed 
in Sec. 29.397, that part of the system must be designed to withstand 
the maximum loads that can be obtained in normal operation. The minimum 
design loads must, in any case, provide a rugged system for service use, 
including consideration of fatigue, jamming, ground gusts, control 
inertia, and friction loads. In the absence of a rational analysis, the 
design loads resulting from 0.60 of the specified limit pilot forces are 
acceptable minimum design loads; and
    (4) If operational loads may be exceeded through jamming, ground 
gusts, control inertia, or friction, the system must withstand the limit 
pilot forces specified in Sec. 29.397, without yielding.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 55 
FR 8002, Mar. 6, 1990]



Sec. 29.397  Limit pilot forces and torques.

    (a) Except as provided in paragraph (b) of this section, the limit 
pilot forces are as follows:
    (1) For foot controls, 130 pounds.
    (2) For stick controls, 100 pounds fore and aft, and 67 pounds 
laterally.

[[Page 761]]

    (b) For flap, tab, stabilizer, rotor brake, and landing gear 
operating controls, the following apply (R=radius in inches):
    (1) Crank wheel, and lever controls, [1 + R]/3 x 50 pounds, but not 
less than 50 pounds nor more than 100 pounds for hand operated controls 
or 130 pounds for foot operated controls, applied at any angle within 20 
degrees of the plane of motion of the control.
    (2) Twist controls, 80R inch-pounds.

[Amdt. 29-12, 41 FR 55471, Dec. 20, 1976, as amended by Amdt. 29-47, 66 
FR 23538, May 9, 2001]



Sec. 29.399  Dual control system.

    Each dual primary flight control system must be able to withstand 
the loads that result when pilot forces not less than 0.75 times those 
obtained under Sec. 29.395 are applied--
    (a) In opposition; and
    (b) In the same direction.



Sec. 29.411  Ground clearance: tail rotor guard.

    (a) It must be impossible for the tail rotor to contact the landing 
surface during a normal landing.
    (b) If a tail rotor guard is required to show compliance with 
paragraph (a) of this section--
    (1) Suitable design loads must be established for the guard: and
    (2) The guard and its supporting structure must be designed to 
withstand those loads.



Sec. 29.427  Unsymmetrical loads.

    (a) Horizontal tail surfaces and their supporting structure must be 
designed for unsymmetrical loads arising from yawing and rotor wake 
effects in combination with the prescribed flight conditions.
    (b) To meet the design criteria of paragraph (a) of this section, in 
the absence of more rational data, both of the following must be met:
    (1) One hundred percent of the maximum loading from the symmetrical 
flight conditions acts on the surface on one side of the plane of 
symmetry, and no loading acts on the other side.
    (2) Fifty percent of the maximum loading from the symmetrical flight 
conditions acts on the surface on each side of the plane of symmetry, in 
opposite directions.
    (c) For empennage arrangements where the horizontal tail surfaces 
are supported by the vertical tail surfaces, the vertical tail surfaces 
and supporting structure must be designed for the combined vertical and 
horizontal surface loads resulting from each prescribed flight 
condition, considered separately. The flight conditions must be selected 
so that the maximum design loads are obtained on each surface. In the 
absence of more rational data, the unsymmetrical horizontal tail surface 
loading distributions described in this section must be assumed.

[Amdt. 27-26, 55 FR 8002, Mar. 6, 1990, as amended by Amdt. 29-31, 55 FR 
38966, Sept. 21, 1990]

                              Ground Loads



Sec. 29.471  General.

    (a) Loads and equilibrium. For limit ground loads--
    (1) The limit ground loads obtained in the landing conditions in 
this part must be considered to be external loads that would occur in 
the rotorcraft structure if it were acting as a rigid body; and
    (2) In each specified landing condition, the external loads must be 
placed in equilibrium with linear and angular inertia loads in a 
rational or conservative manner.
    (b) Critical centers of gravity. The critical centers of gravity 
within the range for which certification is requested must be selected 
so that the maximum design loads are obtained in each landing gear 
element.



Sec. 29.473  Ground loading conditions and assumptions.

    (a) For specified landing conditions, a design maximum weight must 
be used that is not less than the maximum weight. A rotor lift may be 
assumed to act through the center of gravity throughout the landing 
impact. This lift may not exceed two-thirds of the design maximum 
weight.
    (b) Unless otherwise prescribed, for each specified landing 
condition, the rotorcraft must be designed for a limit load factor of 
not less than the limit

[[Page 762]]

inertia load factor substantiated under Sec. 29.725.
    (c) Triggering or actuating devices for additional or supplementary 
energy absorption may not fail under loads established in the tests 
prescribed in Sec. Sec. 29.725 and 29.727, but the factor of safety 
prescribed in Sec. 29.303 need not be used.

[Amdt. 29-3, 33 FR 966, Jan. 26, 1968]



Sec. 29.475  Tires and shock absorbers.

    Unless otherwise prescribed, for each specified landing condition, 
the tires must be assumed to be in their static position and the shock 
absorbers to be in their most critical position.



Sec. 29.477  Landing gear arrangement.

    Sections 29.235, 29.479 through 29.485, and 29.493 apply to landing 
gear with two wheels aft, and one or more wheels forward, of the center 
of gravity.



Sec. 29.479  Level landing conditions.

    (a) Attitudes. Under each of the loading conditions prescribed in 
paragraph (b) of this section, the rotorcraft is assumed to be in each 
of the following level landing attitudes:
    (1) An attitude in which each wheel contacts the ground 
simultaneously.
    (2) An attitude in which the aft wheels contact the ground with the 
forward wheels just clear of the ground.
    (b) Loading conditions. The rotorcraft must be designed for the 
following landing loading conditions:
    (1) Vertical loads applied under Sec. 29.471.
    (2) The loads resulting from a combination of the loads applied 
under paragraph (b)(1) of this section with drag loads at each wheel of 
not less than 25 percent of the vertical load at that wheel.
    (3) The vertical load at the instant of peak drag load combined with 
a drag component simulating the forces required to accelerate the wheel 
rolling assembly up to the specified ground speed, with--
    (i) The ground speed for determination of the spin-up loads being at 
least 75 percent of the optimum forward flight speed for minimum rate of 
descent in autorotation; and
    (ii) The loading conditions of paragraph (b) applied to the landing 
gear and its attaching structure only.
    (4) If there are two wheels forward, a distribution of the loads 
applied to those wheels under paragraphs (b)(1) and (2) of this section 
in a ratio of 40:60.
    (c) Pitching moments. Pitching moments are assumed to be resisted 
by--
    (1) In the case of the attitude in paragraph (a)(1) of this section, 
the forward landing gear; and
    (2) In the case of the attitude in paragraph (a)(2) of this section, 
the angular inertia forces.



Sec. 29.481  Tail-down landing conditions.

    (a) The rotorcraft is assumed to be in the maximum nose-up attitude 
allowing ground clearance by each part of the rotorcraft.
    (b) In this attitude, ground loads are assumed to act perpendicular 
to the ground.



Sec. 29.483  One-wheel landing conditions.

    For the one-wheel landing condition, the rotorcraft is assumed to be 
in the level attitude and to contact the ground on one aft wheel. In 
this attitude--
    (a) The vertical load must be the same as that obtained on that side 
under Sec. 29.479(b)(1); and
    (b) The unbalanced external loads must be reacted by rotorcraft 
inertia.



Sec. 29.485  Lateral drift landing conditions.

    (a) The rotorcraft is assumed to be in the level landing attitude, 
with--
    (1) Side loads combined with one-half of the maximum ground 
reactions obtained in the level landing conditions of Sec. 
29.479(b)(1); and
    (2) The loads obtained under paragraph (a)(1) of this section 
applied--
    (i) At the ground contact point; or
    (ii) For full-swiveling gear, at the center of the axle.
    (b) The rotorcraft must be designed to withstand, at ground 
contact--
    (1) When only the aft wheels contact the ground, side loads of 0.8 
times the vertical reaction acting inward on one side and 0.6 times the 
vertical reaction acting outward on the other side, all

[[Page 763]]

combined with the vertical loads specified in paragraph (a) of this 
section; and
    (2) When the wheels contact the ground simultaneously--
    (i) For the aft wheels, the side loads specified in paragraph (b)(1) 
of this section; and
    (ii) For the forward wheels, a side load of 0.8 times the vertical 
reaction combined with the vertical load specified in paragraph (a) of 
this section.



Sec. 29.493  Braked roll conditions.

    Under braked roll conditions with the shock absorbers in their 
static positions--
    (a) The limit vertical load must be based on a load factor of at 
least--
    (1) 1.33, for the attitude specified in Sec. 29.479(a)(1); and
    (2) 1.0, for the attitude specified in Sec. 29.479(a)(2); and
    (b) The structure must be designed to withstand, at the ground 
contact point of each wheel with brakes, a drag load of at least the 
lesser of--
    (1) The vertical load multiplied by a coefficient of friction of 
0.8; and
    (2) The maximum value based on limiting brake torque.



Sec. 29.497  Ground loading conditions: landing gear with tail wheels.

    (a) General. Rotorcraft with landing gear with two wheels forward 
and one wheel aft of the center of gravity must be designed for loading 
conditions as prescribed in this section.
    (b) Level landing attitude with only the forward wheels contacting 
the ground. In this attitude--
    (1) The vertical loads must be applied under Sec. Sec. 29.471 
through 29.475;
    (2) The vertical load at each axle must be combined with a drag load 
at that axle of not less than 25 percent of that vertical load; and
    (3) Unbalanced pitching moments are assumed to be resisted by 
angular inertia forces.
    (c) Level landing attitude with all wheels contacting the ground 
simultaneously. In this attitude, the rotorcraft must be designed for 
landing loading conditions as prescribed in paragraph (b) of this 
section.
    (d) Maximum nose-up attitude with only the rear wheel contacting the 
ground. The attitude for this condition must be the maximum nose-up 
attitude expected in normal operation, including autorotative landings. 
In this attitude--
    (1) The appropriate ground loads specified in paragraph (b)(1) and 
(2) of this section must be determined and applied, using a rational 
method to account for the moment arm between the rear wheel ground 
reaction and the rotorcraft center of gravity; or
    (2) The probability of landing with initial contact on the rear 
wheel must be shown to be extremely remote.
    (e) Level landing attitude with only one forward wheel contacting 
the ground. In this attitude, the rotorcraft must be designed for ground 
loads as specified in paragraph (b)(1) and (3) of this section.
    (f) Side loads in the level landing attitude. In the attitudes 
specified in paragraphs (b) and (c) of this section, the following 
apply:
    (1) The side loads must be combined at each wheel with one-half of 
the maximum vertical ground reactions obtained for that wheel under 
paragraphs (b) and (c) of this section. In this condition, the side 
loads must be--
    (i) For the forward wheels, 0.8 times the vertical reaction (on one 
side) acting inward, and 0.6 times the vertical reaction (on the other 
side) acting outward; and
    (ii) For the rear wheel, 0.8 times the vertical reaction.
    (2) The loads specified in paragraph (f)(1) of this section must be 
applied--
    (i) At the ground contact point with the wheel in the trailing 
position (for non-full swiveling landing gear or for full swiveling 
landing gear with a lock, steering device, or shimmy damper to keep the 
wheel in the trailing position); or
    (ii) At the center of the axle (for full swiveling landing gear 
without a lock, steering device, or shimmy damper).
    (g) Braked roll conditions in the level landing attitude. In the 
attitudes specified in paragraphs (b) and (c) of this section, and with 
the shock absorbers in their static positions, the rotorcraft must be 
designed for braked roll loads as follows:

[[Page 764]]

    (1) The limit vertical load must be based on a limit vertical load 
factor of not less than--
    (i) 1.0, for the attitude specified in paragraph (b) of this 
section; and
    (ii) 1.33, for the attitude specified in paragraph (c) of this 
section.
    (2) For each wheel with brakes, a drag load must be applied, at the 
ground contact point, of not less than the lesser of--
    (i) 0.8 times the vertical load; and
    (ii) The maximum based on limiting brake torque.
    (h) Rear wheel turning loads in the static ground attitude. In the 
static ground attitude, and with the shock absorbers and tires in their 
static positions, the rotorcraft must be designed for rear wheel turning 
loads as follows:
    (1) A vertical ground reaction equal to the static load on the rear 
wheel must be combined with an equal side load.
    (2) The load specified in paragraph (h)(1) of this section must be 
applied to the rear landing gear--
    (i) Through the axle, if there is a swivel (the rear wheel being 
assumed to be swiveled 90 degrees to the longitudinal axis of the 
rotorcraft); or
    (ii) At the ground contact point if there is a lock, steering device 
or shimmy damper (the rear wheel being assumed to be in the trailing 
position).
    (i) Taxiing condition. The rotorcraft and its landing gear must be 
designed for the loads that would occur when the rotorcraft is taxied 
over the roughest ground that may reasonably be expected in normal 
operation.



Sec. 29.501  Ground loading conditions: landing gear with skids.

    (a) General. Rotorcraft with landing gear with skids must be 
designed for the loading conditions specified in this section. In 
showing compliance with this section, the following apply:
    (1) The design maximum weight, center of gravity, and load factor 
must be determined under Sec. Sec. 29.471 through 29.475.
    (2) Structural yielding of elastic spring members under limit loads 
is acceptable.
    (3) Design ultimate loads for elastic spring members need not exceed 
those obtained in a drop test of the gear with--
    (i) A drop height of 1.5 times that specified in Sec. 29.725; and
    (ii) An assumed rotor lift of not more than 1.5 times that used in 
the limit drop tests prescribed in Sec. 29.725.
    (4) Compliance with paragraph (b) through (e) of this section must 
be shown with--
    (i) The gear in its most critically deflected position for the 
landing condition being considered; and
    (ii) The ground reactions rationally distributed along the bottom of 
the skid tube.
    (b) Vertical reactions in the level landing attitude. In the level 
attitude, and with the rotorcraft contacting the ground along the bottom 
of both skids, the vertical reactions must be applied as prescribed in 
paragraph (a) of this section.
    (c) Drag reactions in the level landing attitude. In the level 
attitude, and with the rotorcraft contacting the ground along the bottom 
of both skids, the following apply:
    (1) The vertical reactions must be combined with horizontal drag 
reactions of 50 percent of the vertical reaction applied at the ground.
    (2) The resultant ground loads must equal the vertical load 
specified in paragraph (b) of this section.
    (d) Sideloads in the level landing attitude. In the level attitude, 
and with the rotorcraft contacting the ground along the bottom of both 
skids, the following apply:
    (1) The vertical ground reaction must be--
    (i) Equal to the vertical loads obtained in the condition specified 
in paragraph (b) of this section; and
    (ii) Divided equally among the skids.
    (2) The vertical ground reactions must be combined with a horizontal 
sideload of 25 percent of their value.
    (3) The total sideload must be applied equally between skids and 
along the length of the skids.
    (4) The unbalanced moments are assumed to be resisted by angular 
inertia.
    (5) The skid gear must be investigated for--
    (i) Inward acting sideloads; and
    (ii) Outward acting sideloads.

[[Page 765]]

    (e) One-skid landing loads in the level attitude. In the level 
attitude, and with the rotorcraft contacting the ground along the bottom 
of one skid only, the following apply:
    (1) The vertical load on the ground contact side must be the same as 
that obtained on that side in the condition specified in paragraph (b) 
of this section.
    (2) The unbalanced moments are assumed to be resisted by angular 
inertia.
    (f) Special conditions. In addition to the conditions specified in 
paragraphs (b) and (c) of this section, the rotorcraft must be designed 
for the following ground reactions:
    (1) A ground reaction load acting up and aft at an angle of 45 
degrees to the longitudinal axis of the rotorcraft. This load must be--
    (i) Equal to 1.33 times the maximum weight;
    (ii) Distributed symmetrically among the skids;
    (iii) Concentrated at the forward end of the straight part of the 
skid tube; and
    (iv) Applied only to the forward end of the skid tube and its 
attachment to the rotorcraft.
    (2) With the rotorcraft in the level landing attitude, a vertical 
ground reaction load equal to one-half of the vertical load determined 
under paragraph (b) of this section. This load must be--
    (i) Applied only to the skid tube and its attachment to the 
rotorcraft; and
    (ii) Distributed equally over 33.3 percent of the length between the 
skid tube attachments and centrally located midway between the skid tube 
attachments.

[Amdt. 29-3, 33 FR 966, Jan. 26, 1968; as amended by Amdt. 27-26, 55 FR 
8002, Mar. 6, 1990]



Sec. 29.505  Ski landing conditions.

    If certification for ski operation is requested, the rotorcraft, 
with skis, must be designed to withstand the following loading 
conditions (where P is the maximum static weight on each ski with the 
rotorcraft at design maximum weight, and n is the limit load factor 
determined under Sec. 29.473(b)):
    (a) Up-load conditions in which--
    (1) A vertical load of Pn and a horizontal load of Pn/4 are 
simultaneously applied at the pedestal bearings; and
    (2) A vertical load of 1.33 P is applied at the pedestal bearings.
    (b) A side load condition in which a side load of 0.35 Pn is applied 
at the pedestal bearings in a horizontal plane perpendicular to the 
centerline of the rotorcraft.
    (c) A torque-load condition in which a torque load of 1.33 P (in 
foot-pounds) is applied to the ski about the vertical axis through the 
centerline of the pedestal bearings.



Sec. 29.511  Ground load: unsymmetrical loads on multiple-wheel units.

    (a) In dual-wheel gear units, 60 percent of the total ground 
reaction for the gear unit must be applied to one wheel and 40 percent 
to the other.
    (b) To provide for the case of one deflated tire, 60 percent of the 
specified load for the gear unit must be applied to either wheel except 
that the vertical ground reaction may not be less than the full static 
value.
    (c) In determining the total load on a gear unit, the transverse 
shift in the load centroid, due to unsymmetrical load distribution on 
the wheels, may be neglected.

[Amdt. 29-3, 33 FR 966, Jan. 26, 1968]

                               Water Loads



Sec. 29.519  Hull type rotorcraft: Water-based and amphibian.

    (a) General. For hull type rotorcraft, the structure must be 
designed to withstand the water loading set forth in paragraphs (b), 
(c), and (d) of this section considering the most severe wave heights 
and profiles for which approval is desired. The loads for the landing 
conditions of paragraphs (b) and (c) of this section must be developed 
and distributed along and among the hull and auxiliary floats, if used, 
in a rational and conservative manner, assuming a rotor lift not 
exceeding two-thirds of the rotorcraft weight to act throughout the 
landing impact.
    (b) Vertical landing conditions. The rotorcraft must initially 
contact the most critical wave surface at zero forward speed in likely 
pitch and roll attitudes which result in critical design

[[Page 766]]

loadings. The vertical descent velocity may not be less than 6.5 feet 
per second relative to the mean water surface.
    (c) Forward speed landing conditions. The rotorcraft must contact 
the most critical wave at forward velocities from zero up to 30 knots in 
likely pitch, roll, and yaw attitudes and with a vertical descent 
velocity of not less than 6.5 feet per second relative to the mean water 
surface. A maximum forward velocity of less than 30 knots may be used in 
design if it can be demonstrated that the forward velocity selected 
would not be exceeded in a normal one-engine-out landing.
    (d) Auxiliary float immersion condition. In addition to the loads 
from the landing conditions, the auxiliary float, and its support and 
attaching structure in the hull, must be designed for the load developed 
by a fully immersed float unless it can be shown that full immersion of 
the float is unlikely, in which case the highest likely float buoyancy 
load must be applied that considers loading of the float immersed to 
create restoring moments compensating for upsetting moments caused by 
side wind, asymmetrical rotorcraft loading, water wave action, and 
rotorcraft inertia.

[Amdt. 29-3, 33 FR 966, Jan. 26, 196; as amended by Amdt. 27-26, 55 FR 
8002, Mar. 6, 1990]



Sec. 29.521  Float landing conditions.

    If certification for float operation (including float amphibian 
operation) is requested, the rotorcraft, with floats, must be designed 
to withstand the following loading conditions (where the limit load 
factor is determined under Sec. 29.473(b) or assumed to be equal to 
that determined for wheel landing gear):
    (a) Up-load conditions in which--
    (1) A load is applied so that, with the rotorcraft in the static 
level attitude, the resultant water reaction passes vertically through 
the center of gravity; and
    (2) The vertical load prescribed in paragraph (a)(1) of this section 
is applied simultaneously with an aft component of 0.25 times the 
vertical component
    (b) A side load condition in which--
    (1) A vertical load of 0.75 times the total vertical load specified 
in paragraph (a)(1) of this section is divided equally among the floats; 
and
    (2) For each float, the load share determined under paragraph (b)(1) 
of this section, combined with a total side load of 0.25 times the total 
vertical load specified in paragraph (b)(1) of this section, is applied 
to that float only.

[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]

                       Main Component Requirements



Sec. 29.547  Main and tail rotor structure.

    (a) A rotor is an assembly of rotating components, which includes 
the rotor hub, blades, blade dampers, the pitch control mechanisms, and 
all other parts that rotate with the assembly.
    (b) Each rotor assembly must be designed as prescribed in this 
section and must function safely for the critical flight load and 
operating conditions. A design assessment must be performed, including a 
detailed failure analysis to identify all failures that will prevent 
continued safe flight or safe landing, and must identify the means to 
minimize the likelihood of their occurrence.
    (c) The rotor structure must be designed to withstand the following 
loads prescribed in Sec. Sec. 29.337 through 29.341 and 29.351:
    (1) Critical flight loads.
    (2) Limit loads occurring under normal conditions of autorotation.
    (d) The rotor structure must be designed to withstand loads 
simulating--
    (1) For the rotor blades, hubs, and flapping hinges, the impact 
force of each blade against its stop during ground operation; and
    (2) Any other critical condition expected in normal operation.
    (e) The rotor structure must be designed to withstand the limit 
torque at any rotational speed, including zero.
    In addition:
    (1) The limit torque need not be greater than the torque defined by 
a torque limiting device (where provided), and may not be less than the 
greater of--
    (i) The maximum torque likely to be transmitted to the rotor 
structure, in either direction, by the rotor drive or by sudden 
application of the rotor brake; and

[[Page 767]]

    (ii) For the main rotor, the limit engine torque specified in Sec. 
29.361.
    (2) The limit torque must be equally and rationally distributed to 
the rotor blades.

(Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, 1425)

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-4, 33 
FR 14106, Sept. 18, 1968; Amdt. 29-40, 61 FR 21907, May 10, 1996]



Sec. 29.549  Fuselage and rotor pylon structures.

    (a) Each fuselage and rotor pylon structure must be designed to 
withstand--
    (1) The critical loads prescribed in Sec. Sec. 29.337 through 
29.341, and 29.351;
    (2) The applicable ground loads prescribed in Sec. Sec. 29.235, 
29.471 through 29.485, 29.493, 29.497, 29.505, and 29.521; and
    (3) The loads prescribed in Sec. 29.547 (d)(1) and (e)(1)(i).
    (b) Auxiliary rotor thrust, the torque reaction of each rotor drive 
system, and the balancing air and inertia loads occurring under 
accelerated flight conditions, must be considered.
    (c) Each engine mount and adjacent fuselage structure must be 
designed to withstand the loads occurring under accelerated flight and 
landing conditions, including engine torque.
    (d) [Reserved]
    (e) If approval for the use of 2\1/2\-minute OEI power is requested, 
each engine mount and adjacent structure must be designed to withstand 
the loads resulting from a limit torque equal to 1.25 times the mean 
torque for 2\1/2\-minute OEI power combined with 1g flight loads.

(Secs. 604, 605, 72 Stat. 778, 49 U.S.C. 1424, 1425)

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-4, 33 
FR 14106, Sept. 18, 1968; Amdt. 29-26, 53 FR 34215, Sept. 2, 1988]



Sec. 29.551  Auxiliary lifting surfaces.

    Each auxiliary lifting surface must be designed to withstand--
    (a) The critical flight loads in Sec. Sec. 29.337 through 29.341, 
and 29.351;
    (b) the applicable ground loads in Sec. Sec. 29.235, 29.471 through 
29.485, 29.493, 29.505, and 29.521; and
    (c) Any other critical condition expected in normal operation.

                      Emergency Landing Conditions



Sec. 29.561  General.

    (a) The rotorcraft, although it may be damaged in emergency landing 
conditions on land or water, must be designed as prescribed in this 
section to protect the occupants under those conditions.
    (b) The structure must be designed to give each occupant every 
reasonable chance of escaping serious injury in a crash landing when--
    (1) Proper use is made of seats, belts, and other safety design 
provisions;
    (2) The wheels are retracted (where applicable); and
    (3) Each occupant and each item of mass inside the cabin that could 
injure an occupant is restrained when subjected to the following 
ultimate inertial load factors relative to the surrounding structure:
    (i) Upward--4g.
    (ii) Forward--16g.
    (iii) Sideward--8g.
    (iv) Downward--20g, after the intended displacement of the seat 
device.
    (v) Rearward--1.5g.
    (c) The supporting structure must be designed to restrain under any 
ultimate inertial load factor up to those specified in this paragraph, 
any item of mass above and/or behind the crew and passenger compartment 
that could injure an occupant if it came loose in an emergency landing. 
Items of mass to be considered include, but are not limited to, rotors, 
transmission, and engines. The items of mass must be restrained for the 
following ultimate inertial load factors:
    (1) Upward--1.5g.
    (2) Forward--12g.
    (3) Sideward--6g.
    (4) Downward--12g.
    (5) Rearward--1.5g.
    (d) Any fuselage structure in the area of internal fuel tanks below 
the passenger floor level must be designed to resist the following 
ultimate inertial factors and loads, and to protect the fuel tanks from 
rupture, if rupture is likely when those loads are applied to that area:
    (1) Upward--1.5g.
    (2) Forward--4.0g.
    (3) Sideward--2.0g.

[[Page 768]]

    (4) Downward--4.0g.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-29, 54 
FR 47319, Nov. 13, 1989; Amdt. 29-38, 61 FR 10438, Mar. 13, 1996]



Sec. 29.562  Emergency landing dynamic conditions.

    (a) The rotorcraft, although it may be damaged in a crash landing, 
must be designed to reasonably protect each occupant when--
    (1) The occupant properly uses the seats, safety belts, and shoulder 
harnesses provided in the design; and
    (2) The occupant is exposed to loads equivalent to those resulting 
from the conditions prescribed in this section.
    (b) Each seat type design or other seating device approved for crew 
or passenger occupancy during takeoff and landing must successfully 
complete dynamic tests or be demonstrated by rational analysis based on 
dynamic tests of a similar type seat in accordance with the following 
criteria. The tests must be conducted with an occupant simulated by a 
170-pound anthropomorphic test dummy (ATD), as defined by 49 CFR 572, 
Subpart B, or its equivalent, sitting in the normal upright position.
    (1) A change in downward velocity of not less than 30 feet per 
second when the seat or other seating device is oriented in its nominal 
position with respect to the rotorcraft's reference system, the 
rotorcraft's longitudinal axis is canted upward 60[deg] with respect to 
the impact velocity vector, and the rotorcraft's lateral axis is 
perpendicular to a vertical plane containing the impact velocity vector 
and the rotorcraft's longitudinal axis. Peak floor deceleration must 
occur in not more than 0.031 seconds after impact and must reach a 
minimum of 30g's.
    (2) A change in forward velocity of not less than 42 feet per second 
when the seat or other seating device is oriented in its nominal 
position with respect to the rotorcraft's reference system, the 
rotorcraft's longitudinal axis is yawed 10[deg] either right or left of 
the impact velocity vector (whichever would cause the greatest load on 
the shoulder harness), the rotorcraft's lateral axis is contained in a 
horizontal plane containing the impact velocity vector, and the 
rotorcraft's vertical axis is perpendicular to a horizontal plane 
containing the impact velocity vector. Peak floor deceleration must 
occur in not more than 0.071 seconds after impact and must reach a 
minimum of 18.4g's.
    (3) Where floor rails or floor or sidewall attachment devices are 
used to attach the seating devices to the airframe structure for the 
conditions of this section, the rails or devices must be misaligned with 
respect to each other by at least 10[deg] vertically (i.e., pitch out of 
parallel) and by at least a 10[deg] lateral roll, with the directions 
optional, to account for possible floor warp.
    (c) Compliance with the following must be shown:
    (1) The seating device system must remain intact although it may 
experience separation intended as part of its design.
    (2) The attachment between the seating device and the airframe 
structure must remain intact although the structure may have exceeded 
its limit load.
    (3) The ATD's shoulder harness strap or straps must remain on or in 
the immediate vicinity of the ATD's shoulder during the impact.
    (4) The safety belt must remain on the ATD's pelvis during the 
impact.
    (5) The ATD's head either does not contact any portion of the crew 
or passenger compartment or, if contact is made, the head impact does 
not exceed a head injury criteria (HIC) of 1,000 as determined by this 
equation.
[GRAPHIC] [TIFF OMITTED] TC28SE91.088

Where: a(t) is the resultant acceleration at the center of gravity of 
          the head form expressed as a multiple of g (the acceleration 
          of gravity) and t2 - t1 is the time 
          duration, in seconds, of major head impact, not to exceed 0.05 
          seconds.

    (6) Loads in individual shoulder harness straps must not exceed 
1,750 pounds. If dual straps are used for retaining the upper torso, the 
total harness strap loads must not exceed 2,000 pounds.
    (7) The maximum compressive load measured between the pelvis and the 
lumbar column of the ATD must not exceed 1,500 pounds.

[[Page 769]]

    (d) An alternate approach that achieves an equivalent or greater 
level of occupant protection, as required by this section, must be 
substantiated on a rational basis.

[Amdt. 29-29, 54 FR 47320, Nov. 13, 1989, as amended by Amdt. 29-41, 62 
FR 46173, Aug. 29, 1997]



Sec. 29.563  Structural ditching provisions.

    If certification with ditching provisions is requested, structural 
strength for ditching must meet the requirements of this section and 
Sec. 29.801(e).
    (a) Forward speed landing conditions. The rotorcraft must initially 
contact the most critical wave for reasonably probable water conditions 
at forward velocities from zero up to 30 knots in likely pitch, roll, 
and yaw attitudes. The rotorcraft limit vertical descent velocity may 
not be less than 5 feet per second relative to the mean water surface. 
Rotor lift may be used to act through the center of gravity throughout 
the landing impact. This lift may not exceed two-thirds of the design 
maximum weight. A maximum forward velocity of less than 30 knots may be 
used in design if it can be demonstrated that the forward velocity 
selected would not be exceeded in a normal one-engine-out touchdown.
    (b) Auxiliary or emergency float conditions--(1) Floats fixed or 
deployed before initial water contact. In addition to the landing loads 
in paragraph (a) of this section, each auxiliary or emergency float, or 
its support and attaching structure in the airframe or fuselage, must be 
designed for the load developed by a fully immersed float unless it can 
be shown that full immersion is unlikely. If full immersion is unlikely, 
the highest likely float buoyancy load must be applied. The highest 
likely buoyancy load must include consideration of a partially immersed 
float creating restoring moments to compensate the upsetting moments 
caused by side wind, unsymmetrical rotorcraft loading, water wave 
action, rotorcraft inertia, and probable structural damage and leakage 
considered under Sec. 29.801(d). Maximum roll and pitch angles 
determined from compliance with Sec. 29.801(d) may be used, if 
significant, to determine the extent of immersion of each float. If the 
floats are deployed in flight, appropriate air loads derived from the 
flight limitations with the floats deployed shall be used in 
substantiation of the floats and their attachment to the rotorcraft. For 
this purpose, the design airspeed for limit load is the float deployed 
airspeed operating limit multiplied by 1.11.
    (2) Floats deployed after initial water contact. Each float must be 
designed for full or partial immersion prescribed in paragraph (b)(1) of 
this section. In addition, each float must be designed for combined 
vertical and drag loads using a relative limit speed of 20 knots between 
the rotorcraft and the water. The vertical load may not be less than the 
highest likely buoyancy load determined under paragraph (b)(1) of this 
section.

[Amdt. 27-26, 55 FR 8003, Mar. 6, 1990]

                           Fatigue Evaluation



Sec. 29.571  Fatigue Tolerance Evaluation of Metallic Structure.

    (a) A fatigue tolerance evaluation of each principal structural 
element (PSE) must be performed, and appropriate inspections and 
retirement time or approved equivalent means must be established to 
avoid catastrophic failure during the operational life of the 
rotorcraft. The fatigue tolerance evaluation must consider the effects 
of both fatigue and the damage determined under paragraph (e)(4) of this 
section. Parts to be evaluated include PSEs of the rotors, rotor drive 
systems between the engines and rotor hubs, controls, fuselage, fixed 
and movable control surfaces, engine and transmission mountings, landing 
gear, and their related primary attachments.
    (b) For the purposes of this section, the term--
    (1) Catastrophic failure means an event that could prevent continued 
safe flight and landing.
    (2) Principal structural element (PSE) means a structural element 
that contributes significantly to the carriage of flight or ground 
loads, and the fatigue failure of that structural element could result 
in catastrophic failure of the aircraft.
    (c) The methodology used to establish compliance with this section 
must

[[Page 770]]

be submitted to and approved by the Administrator.
    (d) Considering all rotorcraft structure, structural elements, and 
assemblies, each PSE must be identified.
    (e) Each fatigue tolerance evaluation required by this section must 
include:
    (1) In-flight measurements to determine the fatigue loads or 
stresses for the PSEs identified in paragraph (d) of this section in all 
critical conditions throughout the range of design limitations required 
by Sec. 29.309 (including altitude effects), except that maneuvering 
load factors need not exceed the maximum values expected in operations.
    (2) The loading spectra as severe as those expected in operations 
based on loads or stresses determined under paragraph (e)(1) of this 
section, including external load operations, if applicable, and other 
high frequency power-cycle operations.
    (3) Takeoff, landing, and taxi loads when evaluating the landing 
gear and other affected PSEs.
    (4) For each PSE identified in paragraph (d) of this section, a 
threat assessment which includes a determination of the probable 
locations, types, and sizes of damage, taking into account fatigue, 
environmental effects, intrinsic and discrete flaws, or accidental 
damage that may occur during manufacture or operation.
    (5) A determination of the fatigue tolerance characteristics for the 
PSE with the damage identified in paragraph (e)(4) of this section that 
supports the inspection and retirement times, or other approved 
equivalent means.
    (6) Analyses supported by test evidence and, if available, service 
experience.
    (f) A residual strength determination is required that substantiates 
the maximum damage size assumed in the fatigue tolerance evaluation. In 
determining inspection intervals based on damage growth, the residual 
strength evaluation must show that the remaining structure, after damage 
growth, is able to withstand design limit loads without failure.
    (g) The effect of damage on stiffness, dynamic behavior, loads, and 
functional performance must be considered.
    (h) Based on the requirements of this section, inspections and 
retirement times or approved equivalent means must be established to 
avoid catastrophic failure. The inspections and retirement times or 
approved equivalent means must be included in the Airworthiness 
Limitations Section of the Instructions for Continued Airworthiness 
required by Section 29.1529 and Section A29.4 of Appendix A of this 
part.
    (i) If inspections for any of the damage types identified in 
paragraph (e)(4) of this section cannot be established within the 
limitations of geometry, inspectability, or good design practice, then 
supplemental procedures, in conjunction with the PSE retirement time, 
must be established to minimize the risk of occurrence of these types of 
damage that could result in a catastrophic failure during the 
operational life of the rotorcraft.

[Doc. No. FAA-2009-0413, Amdt. 29-55, 76 FR 75442, Dec. 2, 2011]



Sec. 29.573  Damage Tolerance and Fatigue Evaluation of Composite
Rotorcraft Structures.

    (a) Each applicant must evaluate the composite rotorcraft structure 
under the damage tolerance standards of paragraph (d) of this section 
unless the applicant establishes that a damage tolerance evaluation is 
impractical within the limits of geometry, inspectability, and good 
design practice. If an applicant establishes that it is impractical 
within the limits of geometry, inspectability, and good design practice, 
the applicant must do a fatigue evaluation in accordance with paragraph 
(e) of this section.
    (b) The methodology used to establish compliance with this section 
must be submitted to and approved by the Administrator.
    (c) Definitions:
    (1) Catastrophic failure is an event that could prevent continued 
safe flight and landing.
    (2) Principal Structural Elements (PSEs) are structural elements 
that contribute significantly to the carrying of flight or ground loads, 
the failure of which could result in catastrophic failure of the 
rotorcraft.

[[Page 771]]

    (3) Threat Assessment is an assessment that specifies the locations, 
types, and sizes of damage, considering fatigue, environmental effects, 
intrinsic and discrete flaws, and impact or other accidental damage 
(including the discrete source of the accidental damage) that may occur 
during manufacture or operation.
    (d) Damage Tolerance Evaluation:
    (1) Each applicant must show that catastrophic failure due to static 
and fatigue loads, considering the intrinsic or discrete manufacturing 
defects or accidental damage, is avoided throughout the operational life 
or prescribed inspection intervals of the rotorcraft by performing 
damage tolerance evaluations of the strength of composite PSEs and other 
parts, detail design points, and fabrication techniques. Each applicant 
must account for the effects of material and process variability along 
with environmental conditions in the strength and fatigue evaluations. 
Each applicant must evaluate parts that include PSEs of the airframe, 
main and tail rotor drive systems, main and tail rotor blades and hubs, 
rotor controls, fixed and movable control surfaces, engine and 
transmission mountings, landing gear, other parts, detail design points, 
and fabrication techniques deemed critical by the FAA. Each damage 
tolerance evaluation must include:
    (i) The identification of all PSEs;
    (ii) In-flight and ground measurements for determining the loads or 
stresses for all PSEs for all critical conditions throughout the range 
of limits in Sec. 29.309 (including altitude effects), except that 
maneuvering load factors need not exceed the maximum values expected in 
service;
    (iii) The loading spectra as severe as those expected in service 
based on loads or stresses determined under paragraph (d)(1)(ii) of this 
section, including external load operations, if applicable, and other 
operations including high-torque events;
    (iv) A threat assessment for all PSEs that specifies the locations, 
types, and sizes of damage, considering fatigue, environmental effects, 
intrinsic and discrete flaws, and impact or other accidental damage 
(including the discrete source of the accidental damage) that may occur 
during manufacture or operation; and
    (v) An assessment of the residual strength and fatigue 
characteristics of all PSEs that supports the replacement times and 
inspection intervals established under paragraph (d)(2) of this section.
    (2) Each applicant must establish replacement times, inspections, or 
other procedures for all PSEs to require the repair or replacement of 
damaged parts before a catastrophic failure. These replacement times, 
inspections, or other procedures must be included in the Airworthiness 
Limitations Section of the Instructions for Continued Airworthiness 
required by Sec. 29.1529.
    (i) Replacement times for PSEs must be determined by tests, or by 
analysis supported by tests, and must show that the structure is able to 
withstand the repeated loads of variable magnitude expected in-service. 
In establishing these replacement times, the following items must be 
considered:
    (A) Damage identified in the threat assessment required by paragraph 
(d)(1)(iv) of this section;
    (B) Maximum acceptable manufacturing defects and in-service damage 
(i.e., those that do not lower the residual strength below ultimate 
design loads and those that can be repaired to restore ultimate 
strength); and
    (C) Ultimate load strength capability after applying repeated loads.
    (ii) Inspection intervals for PSEs must be established to reveal any 
damage identified in the threat assessment required by paragraph 
(d)(1)(iv) of this section that may occur from fatigue or other in-
service causes before such damage has grown to the extent that the 
component cannot sustain the required residual strength capability. In 
establishing these inspection intervals, the following items must be 
considered:
    (A) The growth rate, including no-growth, of the damage under the 
repeated loads expected in-service determined by tests or analysis 
supported by tests;
    (B) The required residual strength for the assumed damage 
established after considering the damage type, inspection interval, 
detectability of damage, and the techniques adopted for damage

[[Page 772]]

detection. The minimum required residual strength is limit load; and
    (C) Whether the inspection will detect the damage growth before the 
minimum residual strength is reached and restored to ultimate load 
capability, or whether the component will require replacement.
    (3) Each applicant must consider the effects of damage on stiffness, 
dynamic behavior, loads, and functional performance on all PSEs when 
substantiating the maximum assumed damage size and inspection interval.
    (e) Fatigue Evaluation: If an applicant establishes that the damage 
tolerance evaluation described in paragraph (d) of this section is 
impractical within the limits of geometry, inspectability, or good 
design practice, the applicant must do a fatigue evaluation of the 
particular composite rotorcraft structure and:
    (1) Identify all PSEs considered in the fatigue evaluation;
    (2) Identify the types of damage for all PSEs considered in the 
fatigue evaluation;
    (3) Establish supplemental procedures to minimize the risk of 
catastrophic failure associated with the damages identified in paragraph 
(d) of this section; and
    (4) Include these supplemental procedures in the Airworthiness 
Limitations section of the Instructions for Continued Airworthiness 
required by Sec. 29.1529.

[Doc. No. FAA-2009-0660, Amdt. 29-59, 76 FR 74664, Dec. 1, 2011]



                    Subpart D_Design and Construction

                                 General



Sec. 29.601  Design.

    (a) The rotorcraft may have no design features or details that 
experience has shown to be hazardous or unreliable.
    (b) The suitability of each questionable design detail and part must 
be established by tests.



Sec. 29.602  Critical parts.

    (a) Critical part. A critical part is a part, the failure of which 
could have a catastrophic effect upon the rotocraft, and for which 
critical characterists have been identified which must be controlled to 
ensure the required level of integrity.
    (b) If the type design includes critical parts, a critical parts 
list shall be established. Procedures shall be established to define the 
critical design characteristics, identify processes that affect those 
characteristics, and identify the design change and process change 
controls necessary for showing compliance with the quality assurance 
requirements of part 21 of this chapter.

[Doc. No. 29311, 64 FR 46232, Aug. 24, 1999]



Sec. 29.603  Materials.

    The suitability and durability of materials used for parts, the 
failure of which could adversely affect safety, must--
    (a) Be established on the basis of experience or tests;
    (b) Meet approved specifications that ensure their having the 
strength and other properties assumed in the design data; and
    (c) Take into account the effects of environmental conditions, such 
as temperature and humidity, expected in service.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424), and sec. 6(c), Dept. of 
Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55471, Dec. 20, 1976; Amdt. 29-17, 43 FR 50599, Oct. 30, 1978]



Sec. 29.605  Fabrication methods.

    (a) The methods of fabrication used must produce consistently sound 
structures. If a fabrication process (such as gluing, spot welding, or 
heat-treating) requires close control to reach this objective, the 
process must be performed according to an approved process 
specification.
    (b) Each new aircraft fabrication method must be substantiated by a 
test program.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29-17, 43 
FR 50599, Oct. 30, 1978]

[[Page 773]]



Sec. 29.607  Fasteners.

    (a) Each removable bolt, screw, nut, pin, or other fastener whose 
loss could jeopardize the safe operation of the rotorcraft must 
incorporate two separate locking devices. The fastener and its locking 
devices may not be adversely affected by the environmental conditions 
associated with the particular installation.
    (b) No self-locking nut may be used on any bolt subject to rotation 
in operation unless a nonfriction locking device is used in addition to 
the self-locking device.

[Amdt. 29-5, 33 FR 14533, Sept. 27, 1968]



Sec. 29.609  Protection of structure.

    Each part of the structure must--
    (a) Be suitably protected against deterioration or loss of strength 
in service due to any cause, including--
    (1) Weathering;
    (2) Corrosion; and
    (3) Abrasion; and
    (b) Have provisions for ventilation and drainage where necessary to 
prevent the accumulation of corrosive, flammable, or noxious fluids.



Sec. 29.610  Lightning and static electricity protection.

    (a) The rotorcraft structure must be protected against catastrophic 
effects from lightning.
    (b) For metallic components, compliance with paragraph (a) of this 
section may be shown by--
    (1) Electrically bonding the components properly to the airframe; or
    (2) Designing the components so that a strike will not endanger the 
rotorcraft.
    (c) For nonmetallic components, compliance with paragraph (a) of 
this section may be shown by--
    (1) Designing the components to minimize the effect of a strike; or
    (2) Incorporating acceptable means of diverting the resulting 
electrical current to not endanger the rotorcraft.
    (d) The electric bonding and protection against lightning and static 
electricity must--
    (1) Minimize the accumulation of electrostatic charge;
    (2) Minimize the risk of electric shock to crew, passengers, and 
service and maintenance personnel using normal precautions;
    (3) Provide and electrical return path, under both normal and fault 
conditions, on rotorcraft having grounded electrical systems; and
    (4) Reduce to an acceptable level the effects of static electricity 
on the functioning of essential electrical and electronic equipment.

[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984; Amdt. 29-40, 61 FR 21907, May 
10, 1996; 61 FR 33963, July 1, 1996; Amdt. 29-53, 76 FR 33135, June 8, 
2011]



Sec. 29.611  Inspection provisions.

    There must be means to allow close examination of each part that 
requires--
    (a) Recurring inspection;
    (b) Adjustment for proper alignment and functioning; or
    (c) Lubrication.



Sec. 29.613  Material strength properties and design values.

    (a) Material strength properties must be based on enough tests of 
material meeting specifications to establish design values on a 
statistical basis.
    (b) Design values must be chosen to minimize the probability of 
structural failure due to material variability. Except as provided in 
paragraphs (d) and (e) of this section, compliance with this paragraph 
must be shown by selecting design values that assure material strength 
with the following probability--
    (1) Where applied loads are eventually distributed through a single 
member within an assembly, the failure of which would result in loss of 
structural integrity of the component, 99 percent probability with 95 
percent confidence; and
    (2) For redundant structures, those in which the failure of 
individual elements would result in applied loads being safely 
distributed to other load-carrying members, 90 percent probability with 
95 percent confidence.
    (c) The strength, detail design, and fabrication of the structure 
must minimize the probability of disastrous fatigue failure, 
particularly at points of stress concentration.
    (d) Design values may be those contained in the following 
publications

[[Page 774]]

(available from the Naval Publications and Forms Center, 5801 Tabor 
Avenue, Philadelphia, PA 19120) or other values approved by the 
Administrator:
    (1) MIL--HDBK-5, ``Metallic Materials and Elements for Flight 
Vehicle Structure''.
    (2) MIL--HDBK-17, ``Plastics for Flight Vehicles''.
    (3) ANC-18, ``Design of Wood Aircraft Structures''.
    (4) MIL--HDBK-23, ``Composite Construction for Flight Vehicles''.
    (e) Other design values may be used if a selection of the material 
is made in which a specimen of each individual item is tested before use 
and it is determined that the actual strength properties of that 
particular item will equal or exceed those used in design.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-17, 43 
FR 50599, Oct. 30, 1978; Amdt. 29-30, 55 FR 8003, Mar. 6, 1990]



Sec. 29.619  Special factors.

    (a) The special factors prescribed in Sec. Sec. 29.621 through 
29.625 apply to each part of the structure whose strength is--
    (1) Uncertain;
    (2) Likely to deteriorate in service before normal replacement; or
    (3) Subject to appreciable variability due to--
    (i) Uncertainties in manufacturing processes; or
    (ii) Uncertainties in inspection methods.
    (b) For each part of the rotorcraft to which Sec. Sec. 29.621 
through 29.625 apply, the factor of safety prescribed in Sec. 29.303 
must be multiplied by a special factor equal to--
    (1) The applicable special factors prescribed in Sec. Sec. 29.621 
through 29.625; or
    (2) Any other factor great enough to ensure that the probability of 
the part being understrength because of the uncertainties specified in 
paragraph (a) of this section is extremely remote.



Sec. 29.621  Casting factors.

    (a) General. The factors, tests, and inspections specified in 
paragraphs (b) and (c) of this section must be applied in addition to 
those necessary to establish foundry quality control. The inspections 
must meet approved specifications. Paragraphs (c) and (d) of this 
section apply to structural castings except castings that are pressure 
tested as parts of hydraulic or other fluid systems and do not support 
structural loads.
    (b) Bearing stresses and surfaces. The casting factors specified in 
paragraphs (c) and (d) of this section--
    (1) Need not exceed 1.25 with respect to bearing stresses regardless 
of the method of inspection used; and
    (2) Need not be used with respect to the bearing surfaces of a part 
whose bearing factor is larger than the applicable casting factor.
    (c) Critical castings. For each casting whose failure would preclude 
continued safe flight and landing of the rotorcraft or result in serious 
injury to any occupant, the following apply:
    (1) Each critical casting must--
    (i) Have a casting factor of not less than 1.25; and
    (ii) Receive 100 percent inspection by visual, radiographic, and 
magnetic particle (for ferromagnetic materials) or penetrant (for 
nonferromagnetic materials) inspection methods or approved equivalent 
inspection methods.
    (2) For each critical casting with a casting factor less than 1.50, 
three sample castings must be static tested and shown to meet--
    (i) The strength requirements of Sec. 29.305 at an ultimate load 
corresponding to a casting factor of 1.25; and
    (ii) The deformation requirements of Sec. 29.305 at a load of 1.15 
times the limit load.
    (d) Noncritical castings. For each casting other than those 
specified in paragraph (c) of this section, the following apply:
    (1) Except as provided in paragraphs (d)(2) and (3) of this section, 
the casting factors and corresponding inspections must meet the 
following table:

------------------------------------------------------------------------
              Casting factor                         Inspection
------------------------------------------------------------------------
2.0 or greater...........................  100 percent visual.
Less than 2.0, greater than 1.5..........  100 percent visual, and
                                            magnetic particle
                                            (ferromagnetic materials),
                                            penetrant (nonferromagnetic
                                            materials), or approved
                                            equivalent inspection
                                            methods.

[[Page 775]]

 
1.25 through 1.50........................  100 percent visual, and
                                            magnetic particle
                                            (ferromagnetic materials),
                                            penetrant (nonferromagnetic
                                            materials), and radiographic
                                            or approved equivalent
                                            inspection methods.
------------------------------------------------------------------------

    (2) The percentage of castings inspected by nonvisual methods may be 
reduced below that specified in paragraph (d)(1) of this section when an 
approved quality control procedure is established.
    (3) For castings procured to a specification that guarantees the 
mechanical properties of the material in the casting and provides for 
demonstration of these properties by test of coupons cut from the 
castings on a sampling basis--
    (i) A casting factor of 1.0 may be used; and
    (ii) The castings must be inspected as provided in paragraph (d)(1) 
of this section for casting factors of ``1.25 through 1.50'' and tested 
under paragraph (c)(2) of this section.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-41, 62 
FR 46173, Aug. 29, 1997]



Sec. 29.623  Bearing factors.

    (a) Except as provided in paragraph (b) of this section, each part 
that has clearance (free fit), and that is subject to pounding or 
vibration, must have a bearing factor large enough to provide for the 
effects of normal relative motion.
    (b) No bearing factor need be used on a part for which any larger 
special factor is prescribed.



Sec. 29.625  Fitting factors.

    For each fitting (part or terminal used to join one structural 
member to another) the following apply:
    (a) For each fitting whose strength is not proven by limit and 
ultimate load tests in which actual stress conditions are simulated in 
the fitting and surrounding structures, a fitting factor of at least 
1.15 must be applied to each part of--
    (1) The fitting;
    (2) The means of attachment; and
    (3) The bearing on the joined members.
    (b) No fitting factor need be used--
    (1) For joints made under approved practices and based on 
comprehensive test data (such as continuous joints in metal plating, 
welded joints, and scarf joints in wood); and
    (2) With respect to any bearing surface for which a larger special 
factor is used.
    (c) For each integral fitting, the part must be treated as a fitting 
up to the point at which the section properties become typical of the 
member.
    (d) Each seat, berth, litter, safety belt, and harness attachment to 
the structure must be shown by analysis, tests, or both, to be able to 
withstand the inertia forces prescribed in Sec. 29.561(b)(3) multiplied 
by a fitting factor of 1.33.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-42, 63 
FR 43285, Aug. 12, 1998]



Sec. 29.629  Flutter and divergence.

    Each aerodynamic surface of the rotorcraft must be free from flutter 
and divergence under each appropriate speed and power condition.

[Doc. No. 28008, 61 FR 21907, May 10, 1996]



Sec. 29.631  Bird strike.

    The rotorcraft must be designed to ensure capability of continued 
safe flight and landing (for Category A) or safe landing (for Category 
B) after impact with a 2.2-lb (1.0 kg) bird when the velocity of the 
rotorcraft (relative to the bird along the flight path of the 
rotorcraft) is equal to VNE or VH (whichever is 
the lesser) at altitudes up to 8,000 feet. Compliance must be shown by 
tests or by analysis based on tests carried out on sufficiently 
representative structures of similar design.

[Doc. No. 28008, 61 FR 21907, May 10, 1996; 61 FR 33963, July 1, 1996]

                                 Rotors



Sec. 29.653  Pressure venting and drainage of rotor blades.

    (a) For each rotor blade--
    (1) There must be means for venting the internal pressure of the 
blade;
    (2) Drainage holes must be provided for the blade; and

[[Page 776]]

    (3) The blade must be designed to prevent water from becoming 
trapped in it.
    (b) Paragraphs (a)(1) and (2) of this section does not apply to 
sealed rotor blades capable of withstanding the maximum pressure 
differentials expected in service.

[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]



Sec. 29.659  Mass balance.

    (a) The rotor and blades must be mass balanced as necessary to--
    (1) Prevent excessive vibration; and
    (2) Prevent flutter at any speed up to the maximum forward speed.
    (b) The structural integrity of the mass balance installation must 
be substantiated.

[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]



Sec. 29.661  Rotor blade clearance.

    There must be enough clearance between the rotor blades and other 
parts of the structure to prevent the blades from striking any part of 
the structure during any operating condition.

[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]



Sec. 29.663  Ground resonance prevention means.

    (a) The reliability of the means for preventing ground resonance 
must be shown either by analysis and tests, or reliable service 
experience, or by showing through analysis or tests that malfunction or 
failure of a single means will not cause ground resonance.
    (b) The probable range of variations, during service, of the damping 
action of the ground resonance prevention means must be established and 
must be investigated during the test required by Sec. 29.241.

[Amdt. 27-26, 55 FR 8003, Mar. 6, 1990]

                             Control Systems



Sec. 29.671  General.

    (a) Each control and control system must operate with the ease, 
smoothness, and positiveness appropriate to its function.
    (b) Each element of each flight control system must be designed, or 
distinctively and permanently marked, to minimize the probability of any 
incorrect assembly that could result in the malfunction of the system.
    (c) A means must be provided to allow full control movement of all 
primary flight controls prior to flight, or a means must be provided 
that will allow the pilot to determine that full control authority is 
available prior to flight.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44437, Nov. 6, 1984]



Sec. 29.672  Stability augmentation, automatic, and power-operated
systems.

    If the functioning of stability augmentation or other automatic or 
power-operated system is necessary to show compliance with the flight 
characteristics requirements of this part, the system must comply with 
Sec. 29.671 of this part and the following:
    (a) A warning which is clearly distinguishable to the pilot under 
expected flight conditions without requiring the pilot's attention must 
be provided for any failure in the stability augmentation system or in 
any other automatic or power-operated system which could result in an 
unsafe condition if the pilot is unaware of the failure. Warning systems 
must not activate the control systems.
    (b) The design of the stability augmentation system or of any other 
automatic or power-operated system must allow initial counteraction of 
failures without requiring exceptional pilot skill or strength, by 
overriding the failure by moving the flight controls in the normal 
sense, and by deactivating the failed system.
    (c) It must be show that after any single failure of the stability 
augmentation system or any other automatic or power-operated system--
    (1) The rotorcraft is safely controllable when the failure or 
malfunction occurs at any speed or altitude within the approved 
operating limitations;
    (2) The controllability and maneuverability requirements of this 
part are met within a practical operational flight envelope (for 
example, speed, altitude, normal acceleration, and rotorcraft 
configurations) which is described in the Rotorcraft Flight Manual; and
    (3) The trim and stability characteristics are not impaired below a 
level

[[Page 777]]

needed to allow continued safe flight and landing.

[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]



Sec. 29.673  Primary flight controls.

    Primary flight controls are those used by the pilot for immediate 
control of pitch, roll, yaw, and vertical motion of the rotorcraft.

[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]



Sec. 29.674  Interconnected controls.

    Each primary flight control system must provide for safe flight and 
landing and operate independently after a malfunction, failure, or jam 
of any auxiliary interconnected control.

[Amdt. 27-26, 55 FR 8003, Mar. 6, 1990]



Sec. 29.675  Stops.

    (a) Each control system must have stops that positively limit the 
range of motionof the pilot's controls.
    (b) Each stop must be located in the system so that the range of 
travel of its control is not appreciably affected by--
    (1) Wear;
    (2) Slackness; or
    (3) Takeup adjustments.
    (c) Each stop must be able to withstand the loads corresponding to 
the design conditions for the system.
    (d) For each main rotor blade--
    (1) Stops that are appropriate to the blade design must be provided 
to limit travel of the blade about its hinge points; and
    (2) There must be means to keep the blade from hitting the droop 
stops during any operation other than starting and stopping the rotor.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29-17, 43 
FR 50599, Oct. 30, 1978]



Sec. 29.679  Control system locks.

    If there is a device to lock the control system with the rotorcraft 
on the ground or water, there must be means to--
    (a) Automatically disengage the lock when the pilot operates the 
controls in a normal manner, or limit the operation of the rotorcraft so 
as to give unmistakable warning to the pilot before takeoff; and
    (b) Prevent the lock from engaging in flight.



Sec. 29.681  Limit load static tests.

    (a) Compliance with the limit load requirements of this part must be 
shown by tests in which--
    (1) The direction of the test loads produces the most severe loading 
in the control system; and
    (2) Each fitting, pulley, and bracket used in attaching the system 
to the main structure is included;
    (b) Compliance must be shown (by analyses or individual load tests) 
with the special factor requirements for control system joints subject 
to angular motion.



Sec. 29.683  Operation tests.

    It must be shown by operation tests that, when the controls are 
operated from the pilot compartment with the control system loaded to 
correspond with loads specified for the system, the system is free 
from--
    (a) Jamming;
    (b) Excessive friction; and
    (c) Excessive deflection.



Sec. 29.685  Control system details.

    (a) Each detail of each control system must be designed to prevent 
jamming, chafing, and interference from cargo, passengers, loose 
objects, or the freezing of moisture.
    (b) There must be means in the cockpit to prevent the entry of 
foreign objects into places where they would jam the system.
    (c) There must be means to prevent the slapping of cables or tubes 
against other parts.
    (d) Cable systems must be designed as follows:
    (1) Cables, cable fittings, turnbuckles, splices, and pulleys must 
be of an acceptable kind.
    (2) The design of cable systems must prevent any hazardous change in 
cable tension throughout the range of travel under any operating 
conditions and temperature variations.

[[Page 778]]

    (3) No cable smaller than \1/8\ inch diameter may be used in any 
primary control system.
    (4) Pulley kinds and sizes must correspond to the cables with which 
they are used. The pulley-cable combinations and strength values 
specified in MIL-HDBK-5 must be used unless they are inapplicable.
    (5) Pulleys must have close fitting guards to prevent the cables 
from being displaced or fouled.
    (6) Pulleys must lie close enough to the plane passing through the 
cable to prevent the cable from rubbing against the pulley flange.
    (7) No fairlead may cause a change in cable direction of more than 
three degrees.
    (8) No clevis pin subject to load or motion and retained only by 
cotter pins may be used in the control system.
    (9) Turnbuckles attached to parts having angular motion must be 
installed to prevent binding throughout the range of travel.
    (10) There must be means for visual inspection at each fairlead, 
pulley, terminal, and turnbuckle.
    (e) Control system joints subject to angular motion must incorporate 
the following special factors with respect to the ultimate bearing 
strength of the softest material used as a bearing:
    (1) 3.33 for push-pull systems other than ball and roller bearing 
systems.
    (2) 2.0 for cable systems.
    (f) For control system joints, the manufacturer's static, non-
Brinell rating of ball and roller bearings may not be exceeded.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55471, Dec. 20, 1976]



Sec. 29.687  Spring devices.

    (a) Each control system spring device whose failure could cause 
flutter or other unsafe characteristics must be reliable.
    (b) Compliance with paragraph (a) of this section must be shown by 
tests simulating service conditions.



Sec. 29.691  Autorotation control mechanism.

    Each main rotor blade pitch control mechanism must allow rapid entry 
into autorotation after power failure.



Sec. 29.695  Power boost and power-operated control system.

    (a) If a power boost or power-operated control system is used, an 
alternate system must be immediately available that allows continued 
safe flight and landing in the event of--
    (1) Any single failure in the power portion of the system; or
    (2) The failure of all engines.
    (b) Each alternate system may be a duplicate power portion or a 
manually operated mechanical system. The power portion includes the 
power source (such as hydrualic pumps), and such items as valves, lines, 
and actuators.
    (c) The failure of mechanical parts (such as piston rods and links), 
and the jamming of power cylinders, must be considered unless they are 
extremely improbable.

                              Landing Gear



Sec. 29.723  Shock absorption tests.

    The landing inertia load factor and the reserve energy absorption 
capacity of the landing gear must be substantiated by the tests 
prescribed in Sec. Sec. 29.725 and 29.727, respectively. These tests 
must be conducted on the complete rotorcraft or on units consisting of 
wheel, tire, and shock absorber in their proper relation.



Sec. 29.725  Limit drop test.

    The limit drop test must be conducted as follows:
    (a) The drop height must be at least 8 inches.
    (b) If considered, the rotor lift specified in Sec. 29.473(a) must 
be introduced into the drop test by appropriate energy absorbing devices 
or by the use of an effective mass.
    (c) Each landing gear unit must be tested in the attitude simulating 
the landing condition that is most critical from the standpoint of the 
energy to be absorbed by it.
    (d) When an effective mass is used in showing compliance with 
paragraph (b)

[[Page 779]]

of this section, the following formulae may be used instead of more 
rational computations.
[GRAPHIC] [TIFF OMITTED] TC28SE91.089

where:

We=the effective weight to be used in the drop test (lbs.).
W=WM for main gear units (lbs.), equal to the static reaction on the 
          particular unit with the rotorcraft in the most critical 
          attitude. A rational method may be used in computing a main 
          gear static reaction, taking into consideration the moment arm 
          between the main wheel reaction and the rotorcraft center of 
          gravity.
W=WN for nose gear units (lbs.), equal to the vertical component of the 
          static reaction that would exist at the nose wheel, assuming 
          that the mass of the rotorcraft acts at the center of gravity 
          and exerts a force of 1.0g downward and 0.25g forward.
W=Wt for tailwheel units (lbs.) equal to whichever of the following is 
          critical--
    (1) The static weight on the tailwheel with the rotorcraft resting 
on all wheels; or
    (2) The vertical component of the ground reaction that would occur 
at the tailwheel assuming that the mass of the rotorcraft acts at the 
center of gravity and exerts a force of 1g downward with the rotorcraft 
in the maximum nose-up attitude considered in the nose-up landing 
conditions.

h=specified free drop height (inches).
L=ratio of assumed rotor lift to the rotorcraft weight.
d=deflection under impact of the tire (at the proper inflation pressure) 
          plus the vertical component of the axle travel (inches) 
          relative to the drop mass.
n=limit inertia load factor.
nj=the load factor developed, during impact, on the mass used in the 
          drop test (i.e., the acceleration dv/dt in g's recorded in the 
          drop test plus 1.0).

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 967, Jan. 26. 1968]



Sec. 29.727  Reserve energy absorption drop test.

    The reserve energy absorption drop test must be conducted as 
follows:
    (a) The drop height must be 1.5 times that specified in Sec. 
29.725(a).
    (b) Rotor lift, where considered in a manner similar to that 
prescribed in Sec. 29.725(b), may not exceed 1.5 times the lift allowed 
under that paragraph.
    (c) The landing gear must withstand this test without collapsing. 
Collapse of the landing gear occurs when a member of the nose, tail, or 
main gear will not support the rotorcraft in the proper attitude or 
allows the rotorcraft structure, other than landing gear and external 
accessories, to impact the landing surface.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 27-26, 55 
FR 8003, Mar. 6, 1990]



Sec. 29.729  Retracting mechanism.

    For rotorcraft with retractable landing gear, the following apply:
    (a) Loads. The landing gear, retracting mechanism, wheel well doors, 
and supporting structure must be designed for--
    (1) The loads occurring in any maneuvering condition with the gear 
retracted;
    (2) The combined friction, inertia, and air loads occurring during 
retraction and extension at any airspeed up to the design maximum 
landing gear operating speed; and
    (3) The flight loads, including those in yawed flight, occurring 
with the gear extended at any airspeed up to the design maximum landing 
gear extended speed.
    (b) Landing gear lock. A positive means must be provided to keep the 
gear extended.
    (c) Emergency operation. When other than manual power is used to 
operate the gear, emergency means must be provided for extending the 
gear in the event of--
    (1) Any reasonably probable failure in the normal retraction system; 
or
    (2) The failure of any single source of hydraulic, electric, or 
equivalent energy.
    (d) Operation tests. The proper functioning of the retracting 
mechanism must be shown by operation tests.
    (e) Position indicator. There must be means to indicate to the pilot 
when the gear is secured in the extreme positions.
    (f) Control. The location and operation of the retraction control 
must meet the requirements of Sec. Sec. 29.777 and 29.779.
    (g) Landing gear warning. An aural or equally effective landing gear 
warning

[[Page 780]]

device must be provided that functions continuously when the rotorcraft 
is in a normal landing mode and the landing gear is not fully extended 
and locked. A manual shutoff capability must be provided for the warning 
device and the warning system must automatically reset when the 
rotorcraft is no longer in the landing mode.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44437, Nov. 6, 1984]



Sec. 29.731  Wheels.

    (a) Each landing gear wheel must be approved.
    (b) The maximum static load rating of each wheel may not be less 
than the corresponding static ground reaction with--
    (1) Maximum weight; and
    (2) Critical center of gravity.
    (c) The maximum limit load rating of each wheel must equal or exceed 
the maximum radial limit load determined under the applicable ground 
load requirements of this part.



Sec. 29.733  Tires.

    Each landing gear wheel must have a tire--
    (a) That is a proper fit on the rim of the wheel; and
    (b) Of a rating that is not exceeded under--
    (1) The design maximum weight;
    (2) A load on each main wheel tire equal to the static ground 
reaction corresponding to the critical center of gravity; and
    (3) A load on nose wheel tires (to be compared with the dynamic 
rating established for those tires) equal to the reaction obtained at 
the nose wheel, assuming that the mass of the rotorcraft acts as the 
most critical center of gravity and exerts a force of 1.0 g downward and 
0.25 g forward, the reactions being distributed to the nose and main 
wheels according to the principles of statics with the drag reaction at 
the ground applied only at wheels with brakes.
    (c) Each tire installed on a retractable landing gear system must, 
at the maximum size of the tire type expected in service, have a 
clearance to surrounding structure and systems that is adequate to 
prevent contact between the tire and any part of the structure or 
systems.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55471, Dec. 20, 1976]



Sec. 29.735  Brakes.

    For rotorcraft with wheel-type landing gear, a braking device must 
be installed that is--
    (a) Controllable by the pilot;
    (b) Usable during power-off landings; and
    (c) Adequate to--
    (1) Counteract any normal unbalanced torque when starting or 
stopping the rotor; and
    (2) Hold the rotorcraft parked on a 10-degree slope on a dry, smooth 
pavement.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44437, Nov. 6, 1984]



Sec. 29.737  Skis.

    (a) The maximum limit load rating of each ski must equal or exceed 
the maximum limit load determined under the applicable ground load 
requirements of this part.
    (b) There must be a stabilizing means to maintain the ski in an 
appropriate position during flight. This means must have enough strength 
to withstand the maximum aerodynamic and inertia loads on the ski.

                            Floats and Hulls



Sec. 29.751  Main float buoyancy.

    (a) For main floats, the buoyancy necessary to support the maximum 
weight of the rotorcraft in fresh water must be exceeded by--
    (1) 50 percent, for single floats; and
    (2) 60 percent, for multiple floats.
    (b) Each main float must have enough water-tight compartments so 
that, with any single main float compartment flooded, the mainfloats 
will provide a margin of positive stability great enough to minimize the 
probability of capsizing.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 967, Jan. 26, 1968]

[[Page 781]]



Sec. 29.753  Main float design.

    (a) Bag floats. Each bag float must be designed to withstand--
    (1) The maximum pressure differential that might be developed at the 
maximum altitude for which certification with that float is requested; 
and
    (2) The vertical loads prescribed in Sec. 29.521(a), distributed 
along the length of the bag over three-quarters of its projected area.
    (b) Rigid floats. Each rigid float must be able to withstand the 
vertical, horizontal, and side loads prescribed in Sec. 29.521. An 
appropriate load distribution under critical conditions must be used.



Sec. 29.755  Hull buoyancy.

    Water-based and amphibian rotorcraft. The hull and auxiliary floats, 
if used, must have enough watertight compartments so that, with any 
single compartment of the hull or auxiliary floats flooded, the buoyancy 
of the hull and auxiliary floats, and wheel tires if used, provides a 
margin of positive water stability great enough to minimize the 
probability of capsizing the rotorcraft for the worst combination of 
wave heights and surface winds for which approval is desired.

[Amdt. 29-3, 33 FR 967, Jan. 26, 1968; as amended by Amdt. 27-26, 55 FR 
8003, Mar. 6, 1990]



Sec. 29.757  Hull and auxiliary float strength.

    The hull, and auxiliary floats if used, must withstand the water 
loads prescribed by Sec. 29.519 with a rational and conservative 
distribution of local and distributed water pressures over the hull and 
float bottom.

[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]

                   Personnel and Cargo Accommodations



Sec. 29.771  Pilot compartment.

    For each pilot compartment--
    (a) The compartment and its equipment must allow each pilot to 
perform his duties without unreasonable concentration or fatigue;
    (b) If there is provision for a second pilot, the rotorcraft must be 
controllable with equal safety from either pilot position. Flight and 
powerplant controls must be designed to prevent confusion or inadvertent 
operation when the rotorcraft is piloted from either position;
    (c) The vibration and noise characteristics of cockpit appurtenances 
may not interfere with safe operation;
    (d) Inflight leakage of rain or snow that could distract the crew or 
harm the structure must be prevented.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 967, Jan. 26, 1968; Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]



Sec. 29.773  Pilot compartment view.

    (a) Nonprecipitation conditions. For nonprecipitation conditions, 
the following apply:
    (1) Each pilot compartment must be arranged to give the pilots a 
sufficiently extensive, clear, and undistorted view for safe operation.
    (2) Each pilot compartment must be free of glare and reflection that 
could interfere with the pilot's view. If certification for night 
operation is requested, this must be shown by night flight tests.
    (b) Precipitation conditions. For precipitation conditions, the 
following apply:
    (1) Each pilot must have a sufficiently extensive view for safe 
operation--
    (i) In heavy rain at forward speeds up to VH; and
    (ii) In the most severe icing condition for which certification is 
requested.
    (2) The first pilot must have a window that--
    (i) Is openable under the conditions prescribed in paragraph (b)(1) 
of this section; and
    (ii) Provides the view prescribed in that paragraph.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 967, Jan. 26, 1968]



Sec. 29.775  Windshields and windows.

    Windshields and windows must be made of material that will not break 
into dangerous fragments.

[Amdt. 29-31, 55 FR 38966, Sept. 21, 1990]

[[Page 782]]



Sec. 29.777  Cockpit controls.

    Cockpit controls must be--
    (a) Located to provide convenient operation and to prevent confusion 
and inadvertent operation; and
    (b) Located and arranged with respect to the pilot's seats so that 
there is full and unrestricted movement of each control without 
interference from the cockpit structure or the pilot's clothing when 
pilots from 5[foot]2 to 6[foot]0 in height are 
seated.



Sec. 29.779  Motion and effect of cockpit controls.

    Cockpit controls must be designed so that they operate in accordance 
with the following movements and actuation:
    (a) Flight controls, including the collective pitch control, must 
operate with a sense of motion which corresponds to the effect on the 
rotorcraft.
    (b) Twist-grip engine power controls must be designed so that, for 
lefthand operation, the motion of the pilot's hand is clockwise to 
increase power when the hand is viewed from the edge containing the 
index finger. Other engine power controls, excluding the collective 
control, must operate with a forward motion to increase power.
    (c) Normal landing gear controls must operate downward to extend the 
landing gear.

[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]



Sec. 29.783  Doors.

    (a) Each closed cabin must have at least one adequate and easily 
accessible external door.
    (b) Each external door must be located, and appropriate operating 
procedures must be established, to ensure that persons using the door 
will not be endangered by the rotors, propellers, engine intakes, and 
exhausts when the operating procedures are used.
    (c) There must be means for locking crew and external passenger 
doors and for preventing their opening in flight inadvertently or as a 
result of mechanical failure. It must be possible to open external doors 
from inside and outside the cabin with the rotorcraft on the ground even 
though persons may be crowded against the door on the inside of the 
rotorcraft. The means of opening must be simple and obvious and so 
arranged and marked that it can be readily located and operated.
    (d) There must be reasonable provisions to prevent the jamming of 
any external doors in a minor crash as a result of fuselage deformation 
under the following ultimate inertial forces except for cargo or service 
doors not suitable for use as an exit in an emergency:
    (1) Upward--1.5g.
    (2) Forward--4.0g.
    (3) Sideward--2.0g.
    (4) Downward--4.0g.
    (e) There must be means for direct visual inspection of the locking 
mechanism by crewmembers to determine whether the external doors 
(including passenger, crew, service, and cargo doors) are fully locked. 
There must be visual means to signal to appropriate crewmembers when 
normally used external doors are closed and fully locked.
    (f) For outward opening external doors usable for entrance or 
egress, there must be an auxiliary safety latching device to prevent the 
door from opening when the primary latching mechanism fails. If the door 
does not meet the requirements of paragraph (c) of this section with 
this device in place, suitable operating procedures must be established 
to prevent the use of the device during takeoff and landing.
    (g) If an integral stair is installed in a passenger entry door that 
is qualified as a passenger emergency exit, the stair must be designed 
so that under the following conditions the effectiveness of passenger 
emergency egress will not be impaired:
    (1) The door, integral stair, and operating mechanism have been 
subjected to the inertial forces specified in paragraph (d) of this 
section, acting separately relative to the surrounding structure.
    (2) The rotorcraft is in the normal ground attitude and in each of 
the attitudes corresponding to collapse of one or more legs, or primary 
members, as applicable, of the landing gear.
    (h) Nonjettisonable doors used as ditching emergency exits must have 
means to enable them to be secured in the open position and remain 
secure for

[[Page 783]]

emergency egress in sea state conditions prescribed for ditching.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-20, 45 
FR 60178, Sept. 11, 1980; Amdt. 29-29, 54 FR 47320, Nov. 13, 1989; Amdt. 
27-26, 55 FR 8003, Mar. 6, 1990; Amdt. 29-31, 55 FR 38966, Sept. 21, 
1990]



Sec. 29.785  Seats, berths, litters, safety belts, and harnesses.

    (a) Each seat, safety belt, harness, and adjacent part of the 
rotorcraft at each station designated for occupancy during takeoff and 
landing must be free of potentially injurious objects, sharp edges, 
protuberances, and hard surfaces and must be designed so that a person 
making proper use of these facilities will not suffer serious injury in 
an emergency landing as a result of the inertial factors specified in 
Sec. 29.561(b) and dynamic conditions specified in Sec. 29.562.
    (b) Each occupant must be protected from serious head injury by a 
safety belt plus a shoulder harness that will prevent the head from 
contacting any injurious object, except as provided for in Sec. 
29.562(c)(5). A shoulder harness (upper torso restraint), in combination 
with the safety belt, constitutes a torso restraint system as described 
in TSO-C114.
    (c) Each occupant's seat must have a combined safety belt and 
shoulder harness with a single-point release. Each pilot's combined 
safety belt and shoulder harness must allow each pilot when seated with 
safety belt and shoulder harness fastened to perform all functions 
necessary for flight operations. There must be a means to secure belt 
and harness when not in use to prevent interference with the operation 
of the rotorcraft and with rapid egress in an emergency.
    (d) If seat backs do not have a firm handhold, there must be hand 
grips or rails along each aisle to let the occupants steady themselves 
while using the aisle in moderately rough air.
    (e) Each projecting object that would injure persons seated or 
moving about in the rotorcraft in normal flight must be padded.
    (f) Each seat and its supporting structure must be designed for an 
occupant weight of at least 170 pounds, considering the maximum load 
factors, inertial forces, and reactions between the occupant, seat, and 
safety belt or harness corresponding with the applicable flight and 
ground-load conditions, including the emergency landing conditions of 
Sec. 29.561(b). In addition--
    (1) Each pilot seat must be designed for the reactions resulting 
from the application of the pilot forces prescribed in Sec. 29.397; and
    (2) The inertial forces prescribed in Sec. 29.561(b) must be 
multiplied by a factor of 1.33 in determining the strength of the 
attachment of--
    (i) Each seat to the structure; and
    (ii) Each safety belt or harness to the seat or structure.
    (g) When the safety belt and shoulder harness are combined, the 
rated strength of the safety belt and shoulder harness may not be less 
than that corresponding to the inertial forces specified in Sec. 
29.561(b), considering the occupant weight of at least 170 pounds, 
considering the dimensional characteristics of the restraint system 
installation, and using a distribution of at least a 60-percent load to 
the safety belt and at least a 40-percent load to the shoulder harness. 
If the safety belt is capable of being used without the shoulder 
harness, the inertial forces specified must be met by the safety belt 
alone.
    (h) When a headrest is used, the headrest and its supporting 
structure must be designed to resist the inertia forces specified in 
Sec. 29.561, with a 1.33 fitting factor and a head weight of at least 
13 pounds.
    (i) Each seating device system includes the device such as the seat, 
the cushions, the occupant restraint system and attachment devices.
    (j) Each seating device system may use design features such as 
crushing or separation of certain parts of the seat in the design to 
reduce occupant loads for the emergency landing dynamic conditions of 
Sec. 29.562; otherwise, the system must remain intact and must not 
interfere with rapid evacuation of the rotorcraft.
    (k) For purposes of this section, a litter is defined as a device 
designed to carry a nonambulatory person, primarily in a recumbent 
position, into and on the rotorcraft. Each berth or litter must be 
designed to withstand

[[Page 784]]

the load reaction of an occupant weight of at least 170 pounds when the 
occupant is subjected to the forward inertial factors specified in Sec. 
29.561(b). A berth or litter installed within 15[deg] or less of the 
longitudinal axis of the rotorcraft must be provided with a padded end-
board, cloth diaphragm, or equivalent means that can withstand the 
forward load reaction. A berth or litter oriented greater than 15[deg] 
with the longitudinal axis of the rotorcraft must be equipped with 
appropriate restraints, such as straps or safety belts, to withstand the 
forward reaction. In addition--
    (1) The berth or litter must have a restraint system and must not 
have corners or other protuberances likely to cause serious injury to a 
person occupying it during emergency landing conditions; and
    (2) The berth or litter attachment and the occupant restraint system 
attachments to the structure must be designed to withstand the critical 
loads resulting from flight and ground load conditions and from the 
conditions prescribed in Sec. 29.561(b). The fitting factor required by 
Sec. 29.625(d) shall be applied.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44437, Nov. 6, 1984; Amdt. 29-29, 54 FR 47320, Nov. 13, 1989; Amdt. 
29-42, 63 FR 43285, Aug. 12, 1998]



Sec. 29.787  Cargo and baggage compartments.

    (a) Each cargo and baggage compartment must be designed for its 
placarded maximum weight of contents and for the critical load 
distributions at the appropriate maximum load factors corresponding to 
the specified flight and ground load conditions, except the emergency 
landing conditions of Sec. 29.561.
    (b) There must be means to prevent the contents of any compartment 
from becoming a hazard by shifting under the loads specified in 
paragraph (a) of this section.
    (c) Under the emergency landing conditions of Sec. 29.561, cargo 
and baggage compartments must--
    (1) Be positioned so that if the contents break loose they are 
unlikely to cause injury to the occupants or restrict any of the escape 
facilities provided for use after an emergency landing; or
    (2) Have sufficient strength to withstand the conditions specified 
in Sec. 29.561, including the means of restraint and their attachments 
required by paragraph (b) of this section. Sufficient strength must be 
provided for the maximum authorized weight of cargo and baggage at the 
critical loading distribution.
    (d) If cargo compartment lamps are installed, each lamp must be 
installed so as to prevent contact between lamp bulb and cargo.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55472, Dec. 20, 1976; Amdt. 29-31, 55 FR 38966, Sept. 21, 1990]



Sec. 29.801  Ditching.

    (a) If certification with ditching provisions is requested, the 
rotorcraft must meet the requirements of this section and Sec. Sec. 
29.807(d), 29.1411 and 29.1415.
    (b) Each practicable design measure, compatible with the general 
characteristics of the rotorcraft, must be taken to minimize the 
probability that in an emergency landing on water, the behavior of the 
rotorcraft would cause immediate injury to the occupants or would make 
it impossible for them to escape.
    (c) The probable behavior of the rotorcraft in a water landing must 
be investigated by model tests or by comparison with rotorcraft of 
similar configuration for which the ditching characteristics are known. 
Scoops, flaps, projections, and any other factors likely to affect the 
hydrodynamic characteristics of the rotorcraft must be considered.
    (d) It must be shown that, under reasonably probable water 
conditions, the flotation time and trim of the rotorcraft will allow the 
occupants to leave the rotorcraft and enter the liferafts required by 
Sec. 29.1415. If compliance with this provision is shown by bouyancy 
and trim computations, appropriate allowances must be made for probable 
structural damage and leakage. If the rotorcraft has fuel tanks (with 
fuel jettisoning provisions) that can reasonably be expected to 
withstand a ditching without leakage, the jettisonable

[[Page 785]]

volume of fuel may be considered as bouyancy volume.
    (e) Unless the effects of the collapse of external doors and windows 
are accounted for in the investigation of the probable behavior of the 
rotorcraft in a water landing (as prescribed in paragraphs (c) and (d) 
of this section), the external doors and windows must be designed to 
withstand the probable maximum local pressures.

[Amdt. 29-12, 41 FR 55472, Dec. 20, 1976]



Sec. 29.803  Emergency evacuation.

    (a) Each crew and passenger area must have means for rapid 
evacuation in a crash landing, with the landing gear (1) extended and 
(2) retracted, considering the possibility of fire.
    (b) Passenger entrance, crew, and service doors may be considered as 
emergency exits if they meet the requirements of this section and of 
Sec. Sec. 29.805 through 29.815.
    (c) [Reserved]
    (d) Except as provided in paragraph (e) of this section, the 
following categories of rotorcraft must be tested in accordance with the 
requirements of appendix D of this part to demonstrate that the maximum 
seating capacity, including the crewmembers required by the operating 
rules, can be evacuated from the rotorcraft to the ground within 90 
seconds:
    (1) Rotorcraft with a seating capacity of more than 44 passengers.
    (2) Rotorcraft with all of the following:
    (i) Ten or more passengers per passenger exit as determined under 
Sec. 29.807(b).
    (ii) No main aisle, as described in Sec. 29.815, for each row of 
passenger seats.
    (iii) Access to each passenger exit for each passenger by virtue of 
design features of seats, such as folding or break-over seat backs or 
folding seats.
    (e) A combination of analysis and tests may be used to show that the 
rotorcraft is capable of being evacuated within 90 seconds under the 
conditions specified in Sec. 29.803(d) if the Administrator finds that 
the combination of analysis and tests will provide data, with respect to 
the emergency evacuation capability of the rotorcraft, equivalent to 
that which would be obtained by actual demonstration.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 967, Jan. 26, 1968; Amdt. 27-26, 55 FR 8004, Mar. 6, 1990]



Sec. 29.805  Flight crew emergency exits.

    (a) For rotorcraft with passenger emergency exits that are not 
convenient to the flight crew, there must be flight crew emergency 
exits, on both sides of the rotorcraft or as a top hatch, in the flight 
crew area.
    (b) Each flight crew emergency exit must be of sufficient size and 
must be located so as to allow rapid evacuation of the flight crew. This 
must be shown by test.
    (c) Each exit must not be obstructed by water or flotation devices 
after a ditching. This must be shown by test, demonstration, or 
analysis.

[Amdt. 29-3, 33 FR 968, Jan. 26, 1968; as amended by Amdt. 27-26, 55 FR 
8004, Mar. 6, 1990]



Sec. 29.807  Passenger emergency exits.

    (a) Type. For the purpose of this part, the types of passenger 
emergency exit are as follows:
    (1) Type I. This type must have a rectangular opening of not less 
than 24 inches wide by 48 inches high, with corner radii not greater 
than one-third the width of the exit, in the passenger area in the side 
of the fuselage at floor level and as far away as practicable from areas 
that might become potential fire hazards in a crash.
    (2) Type II. This type is the same as Type I, except that the 
opening must be at least 20 inches wide by 44 inches high.
    (3) Type III. This type is the same as Type I, except that--
    (i) The opening must be at least 20 inches wide by 36 inches high; 
and
    (ii) The exits need not be at floor level.
    (4) Type IV. This type must have a rectangular opening of not less 
than 19 inches wide by 26 inches high, with corner radii not greater 
than one-third the width of the exit, in the side of the fuselage with a 
step-up inside the rotorcraft of not more than 29 inches.

Openings with dimensions larger than those specified in this section may 
be

[[Page 786]]

used, regardless of shape, if the base of the opening has a flat surface 
of not less than the specified width.
    (b) Passenger emergency exits; side-of-fuselage. Emergency exits 
must be accessible to the passengers and, except as provided in 
paragraph (d) of this section, must be provided in accordance with the 
following table:

------------------------------------------------------------------------
                                       Emergency exits for each side of
                                                 the fuselage
     Passenger seating capacity     ------------------------------------
                                                         Type
                                      Type I  Type II    III     Type IV
------------------------------------------------------------------------
1 through 10.......................  .......  .......  .......         1
11 through 19......................  .......  .......     1 or         2
20 through 39......................  .......        1  .......         1
40 through 59......................        1  .......  .......         1
60 through 79......................        1  .......     1 or         2
------------------------------------------------------------------------

    (c) Passenger emergency exits; other than side-of-fuselage. In 
addition to the requirements of paragraph (b) of this section--
    (1) There must be enough openings in the top, bottom, or ends of the 
fuselage to allow evacuation with the rotorcraft on its side; or
    (2) The probability of the rotorcraft coming to rest on its side in 
a crash landing must be extremely remote.
    (d) Ditching emergency exits for passengers. If certification with 
ditching provisions is requested, ditching emergency exits must be 
provided in accordance with the following requirements and must be 
proven by test, demonstration, or analysis unless the emergency exits 
required by paragraph (b) of this section already meet these 
requirements.
    (1) For rotorcraft that have a passenger seating configuration, 
excluding pilots seats, of nine seats or less, one exit above the 
waterline in each side of the rotorcraft, meeting at least the 
dimensions of a Type IV exit.
    (2) For rotorcraft that have a passenger seating configuration, 
excluding pilots seats, of 10 seats or more, one exit above the 
waterline in a side of the rotorcraft meeting at least the dimensions of 
a Type III exit, for each unit (or part of a unit) of 35 passenger 
seats, but no less than two such exits in the passenger cabin, with one 
on each side of the rotorcraft. However, where it has been shown through 
analysis, ditching demonstrations, or any other tests found necessary by 
the Administrator, that the evacuation capability of the rotorcraft 
during ditching is improved by the use of larger exits, or by other 
means, the passenger seat to exit ratio may be increased.
    (3) Flotation devices, whether stowed or deployed, may not interfere 
with or obstruct the exits.
    (e) Ramp exits. One Type I exit only, or one Type II exit only, that 
is required in the side of the fuselage under paragraph (b) of this 
section, may be installed instead in the ramp of floor ramp rotorcraft 
if--
    (1) Its installation in the side of the fuselage is impractical; and
    (2) Its installation in the ramp meets Sec. 29.813.
    (f) Tests. The proper functioning of each emergency exit must be 
shown by test.

[Amdt. 29-3, 33 FR 968, Jan. 26, 1968, as amended by Amdt. 29-12, 41 FR 
55472, Dec. 20, 1976; Amdt. 27-26, 55 FR 8004, Mar. 6, 1990]



Sec. 29.809  Emergency exit arrangement.

    (a) Each emergency exit must consist of a movable door or hatch in 
the external walls of the fuselage and must provide an unobstructed 
opening to the outside.
    (b) Each emergency exit must be openable from the inside and from 
the outside.
    (c) The means of opening each emergency exit must be simple and 
obvious and may not require exceptional effort.
    (d) There must be means for locking each emergency exit and for 
preventing opening in flight inadvertently or as a result of mechanical 
failure.
    (e) There must be means to minimize the probability of the jamming 
of any emergency exit in a minor crash landing as a result of fuselage 
deformation under the ultimate inertial forces in Sec. 29.783(d).
    (f) Except as provided in paragraph (h) of this section, each land-
based rotorcraft emergency exit must have an approved slide as stated in 
paragraph (g) of this section, or its equivalent, to assist occupants in 
descending to the ground from each floor level exit and an approved 
rope, or its equivalent, for all other exits, if the exit threshold is 
more that 6 feet above the ground--
    (1) With the rotorcraft on the ground and with the landing gear 
extended;

[[Page 787]]

    (2) With one or more legs or part of the landing gear collapsed, 
broken, or not extended; and
    (3) With the rotorcraft resting on its side, if required by Sec. 
29.803(d).
    (g) The slide for each passenger emergency exit must be a self-
supporting slide or equivalent, and must be designed to meet the 
following requirements:
    (1) It must be automatically deployed, and deployment must begin 
during the interval between the time the exit opening means is actuated 
from inside the rotorcraft and the time the exit is fully opened. 
However, each passenger emergency exit which is also a passenger 
entrance door or a service door must be provided with means to prevent 
deployment of the slide when the exit is opened from either the inside 
or the outside under nonemergency conditions for normal use.
    (2) It must be automatically erected within 10 seconds after 
deployment is begun.
    (3) It must be of such length after full deployment that the lower 
end is self-supporting on the ground and provides safe evacuation of 
occupants to the ground after collapse of one or more legs or part of 
the landing gear.
    (4) It must have the capability, in 25-knot winds directed from the 
most critical angle, to deploy and, with the assistance of only one 
person, to remain usable after full deployment to evacuate occupants 
safely to the ground.
    (5) Each slide installation must be qualified by five consecutive 
deployment and inflation tests conducted (per exit) without failure, and 
at least three tests of each such five-test series must be conducted 
using a single representative sample of the device. The sample devices 
must be deployed and inflated by the system's primary means after being 
subjected to the inertia forces specified in Sec. 29.561(b). If any 
part of the system fails or does not function properly during the 
required tests, the cause of the failure or malfunction must be 
corrected by positive means and after that, the full series of five 
consecutive deployment and inflation tests must be conducted without 
failure.
    (h) For rotorcraft having 30 or fewer passenger seats and having an 
exit threshold more than 6 feet above the ground, a rope or other assist 
means may be used in place of the slide specified in paragraph (f) of 
this section, provided an evacuation demonstration is accomplished as 
prescribed in Sec. 29.803(d) or (e).
    (i) If a rope, with its attachment, is used for compliance with 
paragraph (f), (g), or (h) of this section, it must--
    (1) Withstand a 400-pound static load; and
    (2) Attach to the fuselage structure at or above the top of the 
emergency exit opening, or at another approved location if the stowed 
rope would reduce the pilot's view in flight.

[Amdt. 29-3, 33 FR 968, Jan. 26, 1968, as amended by Amdt. 29-29, 54 FR 
47321, Nov. 13, 1989; Amdt. 27-26, 55 FR 8004, Mar. 6, 1990]



Sec. 29.811  Emergency exit marking.

    (a) Each passenger emergency exit, its means of access, and its 
means of opening must be conspicuously marked for the guidance of 
occupants using the exits in daylight or in the dark. Such markings must 
be designed to remain visible for rotorcraft equipped for overwater 
flights if the rotorcraft is capsized and the cabin is submerged.
    (b) The identity and location of each passenger emergency exit must 
be recognizable from a distance equal to the width of the cabin.
    (c) The location of each passenger emergency exit must be indicated 
by a sign visible to occupants approaching along the main passenger 
aisle. There must be a locating sign--
    (1) Next to or above the aisle near each floor emergency exit, 
except that one sign may serve two exits if both exists can be seen 
readily from that sign; and
    (2) On each bulkhead or divider that prevents fore and aft vision 
along the passenger cabin, to indicate emergency exits beyond and 
obscured by it, except that if this is not possible the sign may be 
placed at another appropriate location.
    (d) Each passenger emergency exit marking and each locating sign 
must have white letters 1 inch high on a red background 2 inches high, 
be self or electrically illuminated, and have a

[[Page 788]]

minimum luminescence (brightness) of at least 160 microlamberts. The 
colors may be reversed if this will increase the emergency illumination 
of the passenger compartment.
    (e) The location of each passenger emergency exit operating handle 
and instructions for opening must be shown--
    (1) For each emergency exit, by a marking on or near the exit that 
is readable from a distance of 30 inches; and
    (2) For each Type I or Type II emergency exit with a locking 
mechanism released by rotary motion of the handle, by--
    (i) A red arrow, with a shaft at least three-fourths inch wide and a 
head twice the width of the shaft, extending along at least 70 degrees 
of arc at a radius approximately equal to three-fourths of the handle 
length; and
    (ii) The word ``open'' in red letters 1 inch high, placed 
horizontally near the head of the arrow.
    (f) Each emergency exit, and its means of opening, must be marked on 
the outside of the rotorcraft. In addition, the following apply:
    (1) There must be a 2-inch colored band outlining each passenger 
emergency exit, except small rotorcraft with a maximum weight of 12,500 
pounds or less may have a 2-inch colored band outlining each exit 
release lever or device of passenger emergency exits which are normally 
used doors.
    (2) Each outside marking, including the band, must have color 
contrast to be readily distinguishable from the surrounding fuselage 
surface. The contrast must be such that, if the reflectance of the 
darker color is 15 percent or less, the reflectance of the lighter color 
must be at least 45 percent. ``Reflectance'' is the ratio of the 
luminous flux reflected by a body to the luminous flux it receives. When 
the reflectance of the darker color is greater than 15 percent, at least 
a 30 percent difference between its reflectance and the reflectance of 
the lighter color must be provided.
    (g) Exits marked as such, though in excess of the required number of 
exits, must meet the requirements for emergency exits of the particular 
type. Emergency exits need only be marked with the word ``Exit.''

[Amdt. 29-3, 33 FR 968, Jan. 26, 1968, as amended by Amdt. 29-24, 49 FR 
44438, Nov. 6, 1984; Amdt. 27-26, 55 FR 8004, Mar. 6, 1990; Amdt. 29-31, 
55 FR 38967, Sept. 21, 1990]



Sec. 29.812  Emergency lighting.

    For transport Category A rotorcraft, the following apply:
    (a) A source of light with its power supply independent of the main 
lighting system must be installed to--
    (1) Illuminate each passenger emergency exit marking and locating 
sign; and
    (2) Provide enough general lighting in the passenger cabin so that 
the average illumination, when measured at 40-inch intervals at seat 
armrest height on the center line of the main passenger aisle, is at 
least 0.05 foot-candle.
    (b) Exterior emergency lighting must be provided at each emergency 
exit. The illumination may not be less than 0.05 foot-candle (measured 
normal to the direction of incident light) for minimum width on the 
ground surface, with landing gear extended, equal to the width of the 
emergency exit where an evacuee is likely to make first contact with the 
ground outside the cabin. The exterior emergency lighting may be 
provided by either interior or exterior sources with light intensity 
measurements made with the emergency exits open.
    (c) Each light required by paragraph (a) or (b) of this section must 
be operable manually from the cockpit station and from a point in the 
passenger compartment that is readily accessible. The cockpit control 
device must have an ``on,'' ``off,'' and ``armed'' position so that when 
turned on at the cockpit or passenger compartment station or when armed 
at the cockpit station, the emergency lights will either illuminate or 
remain illuminated upon interruption of the rotorcraft's normal electric 
power.
    (d) Any means required to assist the occupants in descending to the 
ground must be illuminated so that the erected assist means is visible 
from the rotorcraft.
    (1) The assist means must be provided with an illumination of not 
less than 0.03 foot-candle (measured normal

[[Page 789]]

to the direction of the incident light) at the ground end of the erected 
assist means where an evacuee using the established escape route would 
normally make first contact with the ground, with the rotorcraft in each 
of the attitudes corresponding to the collapse of one or more legs of 
the landing gear.
    (2) If the emergency lighting subsystem illuminating the assist 
means is independent of the rotorcraft's main emergency lighting system, 
it--
    (i) Must automatically be activated when the assist means is 
erected;
    (ii) Must provide the illumination required by paragraph (d)(1); and
    (iii) May not be adversely affected by stowage.
    (e) The energy supply to each emergency lighting unit must provide 
the required level of illumination for at least 10 minutes at the 
critical ambient conditions after an emergency landing.
    (f) If storage batteries are used as the energy supply for the 
emergency lighting system, they may be recharged from the rotorcraft's 
main electrical power system provided the charging circuit is designed 
to preclude inadvertent battery discharge into charging circuit faults.

[Amdt. 29-24, 49 FR 44438, Nov. 6, 1984]



Sec. 29.813  Emergency exit access.

    (a) Each passageway between passenger compartments, and each 
passageway leading to Type I and Type II emergency exits, must be--
    (1) Unobstructed; and
    (2) At least 20 inches wide.
    (b) For each emergency exit covered by Sec. 29.809(f), there must 
be enough space adjacent to that exit to allow a crewmember to assist in 
the evacuation of passengers without reducing the unobstructed width of 
the passageway below that required for that exit.
    (c) There must be access from each aisle to each Type III and Type 
IV exit, and
    (1) For rotorcraft that have a passenger seating configuration, 
excluding pilot seats, of 20 or more, the projected opening of the exit 
provided must not be obstructed by seats, berths, or other protrusions 
(including seatbacks in any position) for a distance from that exit of 
not less than the width of the narrowest passenger seat installed on the 
rotorcraft;
    (2) For rotorcraft that have a passenger seating configuration, 
excluding pilot seats, of 19 or less, there may be minor obstructions in 
the region described in paragraph (c)(1) of this section, if there are 
compensating factors to maintain the effectiveness of the exit.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55472, Dec. 20, 1976]



Sec. 29.815  Main aisle width.

    The main passenger aisle width between seats must equal or exceed 
the values in the following table:

------------------------------------------------------------------------
                                                 Minimum main passenger
                                                       aisle width
                                               -------------------------
          Passenger seating capacity             Less than    25 Inches
                                                 25 inches     and more
                                                 from floor   from floor
                                                  (inches)     (inches)
------------------------------------------------------------------------
10 or less....................................           12           15
11 through 19.................................           12           20
20 or more....................................           15           20
------------------------------------------------------------------------
\1\ A narrower width not less than 9 inches may be approved when
  substantiated by tests found necessary by the Administrator.


[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55472, Dec. 20, 1976]



Sec. 29.831  Ventilation.

    (a) Each passenger and crew compartment must be ventilated, and each 
crew compartment must have enough fresh air (but not less than 10 cu. 
ft. per minute per crewmember) to let crewmembers perform their duties 
without undue discomfort or fatigue.
    (b) Crew and passenger compartment air must be free from harmful or 
hazardous concentrations of gases or vapors.
    (c) The concentration of carbon monoxide may not exceed one part in 
20,000 parts of air during forward flight. If the concentration exceeds 
this value under other conditions, there must be suitable operating 
restrictions.
    (d) There must be means to ensure compliance with paragraphs (b) and 
(c) of this section under any reasonably probable failure of any 
ventilating, heating, or other system or equipment.

[[Page 790]]



Sec. 29.833  Heaters.

    Each combustion heater must be approved.

                             Fire Protection



Sec. 29.851  Fire extinguishers.

    (a) Hand fire extinguishers. For hand fire extinguishers the 
following apply:
    (1) Each hand fire extinguisher must be approved.
    (2) The kinds and quantities of each extinguishing agent used must 
be appropriate to the kinds of fires likely to occur where that agent is 
used.
    (3) Each extinguisher for use in a personnel compartment must be 
designed to minimize the hazard of toxic gas concentrations.
    (b) Built-in fire extinguishers. If a built-in fire extinguishing 
system is required--
    (1) The capacity of each system, in relation to the volume of the 
compartment where used and the ventilation rate, must be adequate for 
any fire likely to occur in that compartment.
    (2) Each system must be installed so that--
    (i) No extinguishing agent likely to enter personnel compartments 
will be present in a quantity that is hazardous to the occupants; and
    (ii) No discharge of the extinguisher can cause structural damage.



Sec. 29.853  Compartment interiors.

    For each compartment to be used by the crew or passengers--
    (a) The materials (including finishes or decorative surfaces applied 
to the materials) must meet the following test criteria as applicable:
    (1) Interior ceiling panels, interior wall panels, partitions, 
galley structure, large cabinet walls, structural flooring, and 
materials used in the construction of stowage compartments (other than 
underseat stowage compartments and compartments for stowing small items 
such as magazines and maps) must be self-extinguishing when tested 
vertically in accordance with the applicable portions of appendix F of 
Part 25 of this chapter, or other approved equivalent methods. The 
average burn length may not exceed 6 inches and the average flame time 
after removal of the flame source may not exceed 15 seconds. Drippings 
from the test specimen may not continue to flame for more than an 
average of 3 seconds after falling.
    (2) Floor covering, textiles (including draperies and upholstery), 
seat cushions, padding, decorative and nondecorative coated fabrics, 
leather, trays and galley furnishings, electrical conduit, thermal and 
acoustical insulation and insulation covering, air ducting, joint and 
edge covering, cargo compartment liners, insulation blankets, cargo 
covers, and transparencies, molded and thermoformed parts, air ducting 
joints, and trim strips (decorative and chafing) that are constructed of 
materials not covered in paragraph (a)(3) of this section, must be self 
extinguishing when tested vertically in accordance with the applicable 
portion of appendix F of Part 25 of this chapter, or other approved 
equivalent methods. The average burn length may not exceed 8 inches and 
the average flame time after removal of the flame source may not exceed 
15 seconds. Drippings from the test specimen may not continue to flame 
for more than an average of 5 seconds after falling.
    (3) Acrylic windows and signs, parts constructed in whole or in part 
of elastometric materials, edge lighted instrument assemblies consisting 
of two or more instruments in a common housing, seat belts, shoulder 
harnesses, and cargo and baggage tiedown equipment, including 
containers, bins, pallets, etc., used in passenger or crew compartments, 
may not have an average burn rate greater than 2.5 inches per minute 
when tested horizontally in accordance with the applicable portions of 
appendix F of Part 25 of this chapter, or other approved equivalent 
methods.
    (4) Except for electrical wire and cable insulation, and for small 
parts (such as knobs, handles, rollers, fasteners, clips, grommets, rub 
strips, pulleys, and small electrical parts) that the Administrator 
finds would not contribute significantly to the propagation of a fire, 
materials in items not specified in paragraphs (a)(1), (a)(2), or (a)(3) 
of this section may not have a burn rate greater than 4 inches per 
minute when tested horizontally in accordance with the applicable 
portions

[[Page 791]]

of appendix F of Part 25 of this chapter, or other approved equivalent 
methods.
    (b) In addition to meeting the requirements of paragraph (a)(2), 
seat cushions, except those on flight crewmember seats, must meet the 
test requirements of Part II of appendix F of Part 25 of this chapter, 
or equivalent.
    (c) If smoking is to be prohibited, there must be a placard so 
stating, and if smoking is to be allowed--
    (1) There must be an adequate number of self-contained, removable 
ashtrays; and
    (2) Where the crew compartment is separated from the passenger 
compartment, there must be at least one illuminated sign (using either 
letters or symbols) notifying all passengers when smoking is prohibited. 
Signs which notify when smoking is prohibited must--
    (i) When illuminated, be legible to each passenger seated in the 
passenger cabin under all probable lighting conditions; and
    (ii) Be so constructed that the crew can turn the illumination on 
and off.
    (d) Each receptacle for towels, paper, or waste must be at least 
fire-resistant and must have means for containing possible fires;
    (e) There must be a hand fire extinguisher for the flight 
crewmembers; and
    (f) At least the following number of hand fire extinguishers must be 
conveniently located in passenger compartments:

------------------------------------------------------------------------
                                                               Fire
                   Passenger capacity                      extinguishers
------------------------------------------------------------------------
7 through 30............................................               1
31 through 60...........................................               2
61 or more..............................................               3
------------------------------------------------------------------------


(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 969, Jan. 26, 1968; Amdt. 29-17, 43 FR 50600, Oct. 30, 1978; Amdt. 
29-18, 45 FR 7756, Feb. 4, 1980; Amdt. 29-23, 49 FR 43200, Oct. 26, 
1984]



Sec. 29.855  Cargo and baggage compartments.

    (a) Each cargo and baggage compartment must be construced of or 
lined with materials in accordance with the following:
    (1) For accessible and inaccessible compartments not occupied by 
passengers or crew, the material must be at least fire resistant.
    (2) Materials must meet the requirements in Sec. 29.853(a)(1), 
(a)(2), and (a)(3) for cargo or baggage compartments in which--
    (i) The presence of a compartment fire would be easily discovered by 
a crewmember while at the crewmember's station;
    (ii) Each part of the compartment is easily accessible in flight;
    (iii) The compartment has a volume of 200 cubic feet or less; and
    (iv) Notwithstanding Sec. 29.1439(a), protective breathing 
equipment is not required.
    (b) No compartment may contain any controls, wiring, lines, 
equipment, or accessories whose damage or failure would affect safe 
operation, unless those items are protected so that--
    (1) They cannot be damaged by the movement of cargo in the 
compartment; and
    (2) Their breakage or failure will not create a fire hazard.
    (c) The design and sealing of inaccessible compartments must be 
adequate to contain compartment fires until a landing and safe 
evacuation can be made.
    (d) Each cargo and baggage compartment that is not sealed so as to 
contain cargo compartment fires completely without endangering the 
safety of a rotorcraft or its occupants must be designed, or must have a 
device, to ensure detection of fires or smoke by a crewmember while at 
his station and to prevent the accumulation of harmful quantities of 
smoke, flame, extinguishing agents, and other noxious gases in any crew 
or passenger compartment. This must be shown in flight.
    (e) For rotorcraft used for the carriage of cargo only, the cabin 
area may be considered a cargo compartment and, in addition to 
paragraphs (a) through (d) of this section, the following apply:
    (1) There must be means to shut off the ventilating airflow to or 
within the compartment. Controls for this purpose

[[Page 792]]

must be accessible to the flight crew in the crew compartment.
    (2) Required crew emergency exits must be accessible under all cargo 
loading conditions.
    (3) Sources of heat within each compartment must be shielded and 
insulated to prevent igniting the cargo.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 969, Jan. 26, 1968; Amdt. 29-24, 49 FR 44438, Nov. 6, 1984; Amdt. 27-
26, 55 FR 8004, Mar. 6, 1990]



Sec. 29.859  Combustion heater fire protection.

    (a) Combustion heater fire zones. The following combustion heater 
fire zones must be protected against fire under the applicable 
provisions of Sec. Sec. 29.1181 through 29.1191, and 29.1195 through 
29.1203:
    (1) The region surrounding any heater, if that region contains any 
flammable fluid system components (including the heater fuel system), 
that could--
    (i) Be damaged by heater malfunctioning; or
    (ii) Allow flammable fluids or vapors to reach the heater in case of 
leakage.
    (2) Each part of any ventilating air passage that--
    (i) Surrounds the combustion chamber; and
    (ii) Would not contain (without damage to other rotorcraft 
components) any fire that may occur within the passage.
    (b) Ventilating air ducts. Each ventilating air duct passing through 
any fire zone must be fireproof. In addition--
    (1) Unless isolation is provided by fireproof valves or by equally 
effective means, the ventilating air duct downstream of each heater must 
be fireproof for a distance great enough to ensure that any fire 
originating in the heater can be contained in the duct; and
    (2) Each part of any ventilating duct passing through any region 
having a flammable fluid system must be so constructed or isolated from 
that system that the malfunctioning of any component of that system 
cannot introduce flammable fluids or vapors into the ventilating 
airstream.
    (c) Combustion air ducts. Each combustion air duct must be fireproof 
for a distance great enough to prevent damage from backfiring or reverse 
flame propagation. In addition--
    (1) No combustion air duct may communicate with the ventilating 
airstream unless flames from backfires or reverse burning cannot enter 
the ventilating airstream under any operating condition, including 
reverse flow or malfunction of the heater or its associated components; 
and
    (2) No combustion air duct may restrict the prompt relief of any 
backfire that, if so restricted, could cause heater failure.
    (d) Heater controls; general. There must be means to prevent the 
hazardous accumulation of water or ice on or in any heater control 
component, control system tubing, or safety control.
    (e) Heater safety controls. For each combustion heater, safety 
control means must be provided as follows:
    (1) Means independent of the components provided for the normal 
continuous control of air temperature, airflow, and fuel flow must be 
provided, for each heater, to automatically shut off the ignition and 
fuel supply of that heater at a point remote from that heater when any 
of the following occurs:
    (i) The heat exchanger temperature exceeds safe limits.
    (ii) The ventilating air temperature exceeds safe limits.
    (iii) The combustion airflow becomes inadequate for safe operation.
    (iv) The ventilating airflow becomes inadequate for safe operation.
    (2) The means of complying with paragraph (e)(1) of this section for 
any individual heater must--
    (i) Be independent of components serving any other heater whose heat 
output is essential for safe operation; and
    (ii) Keep the heater off until restarted by the crew.
    (3) There must be means to warn the crew when any heater whose heat 
output is essential for safe operation has been shut off by the 
automatic means prescribed in paragraph (e)(1) of this section.
    (f) Air intakes. Each combustion and ventilating air intake must be 
where no flammable fluids or vapors can

[[Page 793]]

enter the heater system under any operating condition--
    (1) During normal operation; or
    (2) As a result of the malfunction of any other component.
    (g) Heater exhaust. Each heater exhaust system must meet the 
requirements of Sec. Sec. 29.1121 and 29.1123. In addition--
    (1) Each exhaust shroud must be sealed so that no flammable fluids 
or hazardous quantities of vapors can reach the exhaust systems through 
joints; and
    (2) No exhaust system may restrict the prompt relief of any backfire 
that, if so restricted, could cause heater failure.
    (h) Heater fuel systems. Each heater fuel system must meet the 
powerplant fuel system requirements affecting safe heater operation. 
Each heater fuel system component in the ventilating airstream must be 
protected by shrouds so that no leakage from those components can enter 
the ventilating airstream.
    (i) Drains. There must be means for safe drainage of any fuel that 
might accumulate in the combustion chamber or the heat exchanger. In 
addition--
    (1) Each part of any drain that operates at high temperatures must 
be protected in the same manner as heater exhausts; and
    (2) Each drain must be protected against hazardous ice accumulation 
under any operating condition.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-2, 32 
FR 6914, May 5, 1967]



Sec. 29.861  Fire protection of structure, controls, and other parts.

    Each part of the structure, controls, and the rotor mechanism, and 
other parts essential to controlled landing and (for category A) flight 
that would be affected by powerplant fires must be isolated under Sec. 
29.1191, or must be--
    (a) For category A rotorcraft, fireproof; and
    (b) For Category B rotorcraft, fireproof or protected so that they 
can perform their essential functions for at least 5 minutes under any 
foreseeable powerplant fire conditions.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 27-26, 55 
FR 8005, Mar. 6, 1990]



Sec. 29.863  Flammable fluid fire protection.

    (a) In each area where flammable fluids or vapors might escape by 
leakage of a fluid system, there must be means to minimize the 
probability of ignition of the fluids and vapors, and the resultant 
hazards if ignition does occur.
    (b) Compliance with paragraph (a) of this section must be shown by 
analysis or tests, and the following factors must be considered:
    (1) Possible sources and paths of fluid leakage, and means of 
detecting leakage.
    (2) Flammability characteristics of fluids, including effects of any 
combustible or absorbing materials.
    (3) Possible ignition sources, including electrical faults, 
overheating of equipment, and malfunctioning of protective devices.
    (4) Means available for controlling or extinguishing a fire, such as 
stopping flow of fluids, shutting down equipment, fireproof containment, 
or use of extinguishing agents.
    (5) Ability of rotorcraft components that are critical to safety of 
flight to withstand fire and heat.
    (c) If action by the flight crew is required to prevent or 
counteract a fluid fire (e.g. equipment shutdown or actuation of a fire 
extinguisher), quick acting means must be provided to alert the crew.
    (d) Each area where flammable fluids or vapors might escape by 
leakage of a fluid system must be identified and defined.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Amdt. 29-17, 43 FR 50600, Oct. 30, 1978]

[[Page 794]]

                             External Loads



Sec. 29.865  External loads.

    (a) It must be shown by analysis, test, or both, that the rotorcraft 
external load attaching means for rotorcraft-load combinations to be 
used for nonhuman external cargo applications can withstand a limit 
static load equal to 2.5, or some lower load factor approved under 
Sec. Sec. 29.337 through 29.341, multiplied by the maximum external 
load for which authorization is requested. It must be shown by analysis, 
test, or both that the rotorcraft external load attaching means and 
corresponding personnel carrying device system for rotorcraft-load 
combinations to be used for human external cargo applications can 
withstand a limit static load equal to 3.5 or some lower load factor, 
not less than 2.5, approved under Sec. Sec. 29.337 through 29.341, 
multiplied by the maximum external load for which authorization is 
requested. The load for any rotorcraft-load combination class, for any 
external cargo type, must be applied in the vertical direction. For 
jettisonable external loads of any applicable external cargo type, the 
load must also be applied in any direction making the maximum angle with 
the vertical that can be achieved in service but not less than 30[deg]. 
However, the 30[deg] angle may be reduced to a lesser angle if--
    (1) An operating limitation is established limiting external load 
operations to such angles for which compliance with this paragraph has 
been shown; or
    (2) It is shown that the lesser angle can not be exceeded in 
service.
    (b) The external load attaching means, for jettisonable rotorcraft-
load combinations, must include a quick-release system to enable the 
pilot to release the external load quickly during flight. The quick-
release system must consist of a primary quick release subsystem and a 
backup quick release subsystem that are isolated from one another. The 
quick release system, and the means by which it is controlled, must 
comply with the following:
    (1) A control for the primary quick release subsystem must be 
installed either on one of the pilot's primary controls or in an 
equivalently accessible location and must be designed and located so 
that it may be operated by either the pilot or a crewmember without 
hazardously limiting the ability to control the rotorcraft during an 
emergency situation.
    (2) A control for the backup quick release subsystem, readily 
accessible to either the pilot or another crewmember, must be provided.
    (3) Both the primary and backup quick release subsystems must--
    (i) Be reliable, durable, and function properly with all external 
loads up to and including the maximum external limit load for which 
authorization is requested.
    (ii) Be protected against electromagnetic interference (EMI) from 
external and internal sources and against lightning to prevent 
inadvertent load release.
    (A) The minimum level of protection required for jettisonable 
rotorcraft-load combinations used for nonhuman external cargo is a radio 
frequency field strength of 20 volts per meter.
    (B) The minimum level of protection required for jettisonable 
rotorcraft-load combinations used for human external cargo is a radio 
frequency field strength of 200 volts per meter.
    (iii) Be protected against any failure that could be induced by a 
failure mode of any other electrical or mechanical rotorcraft system.
    (c) For rotorcraft-load combinations to be used for human external 
cargo applications, the rotorcraft must--
    (1) For jettisonable external loads, have a quick-release system 
that meets the requirements of paragraph (b) of this section and that--
    (i) Provides a dual actuation device for the primary quick release 
subsystem, and
    (ii) Provides a separate dual actuation device for the backup quick 
release subsystem;
    (2) Have a reliable, approved personnel carrying device system that 
has the structural capability and personnel safety features essential 
for external occupant safety;
    (3) Have placards and markings at all appropriate locations that 
clearly state the essential system operating instructions and, for the 
personnel carrying

[[Page 795]]

device system, ingress and egress instructions;
    (4) Have equipment to allow direct intercommunication among required 
crewmembers and external occupants;
    (5) Have the appropriate limitations and procedures incorporated in 
the flight manual for conducting human external cargo operations; and
    (6) For human external cargo applications requiring use of Category 
A rotorcraft, have one-engine-inoperative hover performance data and 
procedures in the flight manual for the weights, altitudes, and 
temperatures for which external load approval is requested.
    (d) The critically configured jettisonable external loads must be 
shown by a combination of analysis, ground tests, and flight tests to be 
both transportable and releasable throughout the approved operational 
envelope without hazard to the rotorcraft during normal flight 
conditions. In addition, these external loads--must be shown to be 
releasable without hazard to the rotorcraft during emergency flight 
conditions.
    (e) A placard or marking must be installed next to the external-load 
attaching means clearly stating any operational limitations and the 
maximum authorized external load as demonstrated under Sec. 29.25 and 
this section.
    (f) The fatigue evaluation of Sec. 29.571 of this part does not 
apply to rotorcraft-load combinations to be used for nonhuman external 
cargo except for the failure of critical structural elements that would 
result in a hazard to the rotorcraft. For rotorcraft-load combinations 
to be used for human external cargo, the fatigue evaluation of Sec. 
29.571 of this part applies to the entire quick release and personnel 
carrying device structural systems and their attachments.

[Amdt. 29-12, 41 FR 55472, Dec. 20, 1976, as amended by Amdt. 27-26, 55 
FR 8005, Mar. 6, 1990; Amdt. 29-43, 64 FR 43020, Aug. 6, 1999]

                              Miscellaneous



Sec. 29.871  Leveling marks.

    There must be reference marks for leveling the rotorcraft on the 
ground.



Sec. 29.873  Ballast provisions.

    Ballast provisions must be designed and constructed to prevent 
inadvertent shifting of ballast in flight.



                          Subpart E_Powerplant

                                 General



Sec. 29.901  Installation.

    (a) For the purpose of this part, the powerplant installation 
includes each part of the rotorcraft (other than the main and auxiliary 
rotor structures) that--
    (1) Is necessary for propulsion;
    (2) Affects the control of the major propulsive units; or
    (3) Affects the safety of the major propulsive units between normal 
inspections or overhauls.
    (b) For each powerplant installation--
    (1) The installation must comply with--
    (i) The installation instructions provided under Sec. 33.5 of this 
chapter; and
    (ii) The applicable provisions of this subpart.
    (2) Each component of the installation must be constructed, 
arranged, and installed to ensure its continued safe operation between 
normal inspections or overhauls for the range of temperature and 
altitude for which approval is requested.
    (3) Accessibility must be provided to allow any inspection and 
maintenance necessary for continued airworthiness; and
    (4) Electrical interconnections must be provided to prevent 
differences of potential between major components of the installation 
and the rest of the rotorcraft.
    (5) Axial and radial expansion of turbine engines may not affect the 
safety of the installation.
    (6) Design precautions must be taken to minimize the possibility of 
incorrect assembly of components and equipment essential to safe 
operation of the rotorcraft, except where operation with the incorrect 
assembly can be shown to be extremely improbable.
    (c) For each powerplant and auxiliary power unit installation, it 
must be established that no single failure or malfunction or probable 
combination of

[[Page 796]]

failures will jeopardize the safe operation of the rotorcraft except 
that the failure of structural elements need not be considered if the 
probability of any such failure is extremely remote.
    (d) Each auxiliary power unit installation must meet the applicable 
provisions of this subpart.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 969, Jan. 26, 1968; Amdt. 29-13, 42 FR 15046, Mar. 17, 1977; Amdt. 
29-17, 43 FR 50600, Oct. 30, 1978; Amdt. 29-26, 53 FR 34215, Sept. 2, 
1988; Amdt. 29-36, 60 FR 55776, Nov. 2, 1995]



Sec. 29.903  Engines.

    (a) Engine type certification. Each engine must have an approved 
type certificate. Reciprocating engines for use in helicopters must be 
qualified in accordance with Sec. 33.49(d) of this chapter or be 
otherwise approved for the intended usage.
    (b) Category A; engine isolation. For each category A rotorcraft, 
the powerplants must be arranged and isolated from each other to allow 
operation, in at least one configuration, so that the failure or 
malfunction of any engine, or the failure of any system that can affect 
any engine, will not--
    (1) Prevent the continued safe operation of the remaining engines; 
or
    (2) Require immediate action, other than normal pilot action with 
primary flight controls, by any crewmember to maintain safe operation.
    (c) Category A; control of engine rotation. For each Category A 
rotorcraft, there must be a means for stopping the rotation of any 
engine individually in flight, except that, for turbine engine 
installations, the means for stopping the engine need be provided only 
where necessary for safety. In addition--
    (1) Each component of the engine stopping system that is located on 
the engine side of the firewall, and that might be exposed to fire, must 
be at least fire resistant; or
    (2) Duplicate means must be available for stopping the engine and 
the controls must be where all are not likely to be damaged at the same 
time in case of fire.
    (d) Turbine engine installation. For turbine engine installations--
    (1) Design precautions must be taken to minimize the hazards to the 
rotorcraft in the event of an engine rotor failure; and
    (2) The powerplant systems associated with engine control devices, 
systems, and instrumentation must be designed to give reasonable 
assurance that those engine operating limitations that adversely affect 
engine rotor structural integrity will not be exceeded in service.
    (e) Restart capability. (1) A means to restart any engine in flight 
must be provided.
    (2) Except for the in-flight shutdown of all engines, engine restart 
capability must be demonstrated throughout a flight envelope for the 
rotorcraft.
    (3) Following the in-flight shutdown of all engines, in-flight 
engine restart capability must be provided.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55472, Dec. 20, 1976; Amdt. 29-26, 53 FR 34215, Sept. 2, 1988; Amdt. 
29-31, 55 FR 38967, Sept. 21, 1990; 55 FR 41309, Oct. 10, 1990; Amdt. 
29-36, 60 FR 55776, Nov. 2, 1995]



Sec. 29.907  Engine vibration.

    (a) Each engine must be installed to prevent the harmful vibration 
of any part of the engine or rotorcraft.
    (b) The addition of the rotor and the rotor drive system to the 
engine may not subject the principal rotating parts of the engine to 
excessive vibration stresses. This must be shown by a vibration 
investigation.



Sec. 29.908  Cooling fans.

    For cooling fans that are a part of a powerplant installation the 
following apply:
    (a) Category A. For cooling fans installed in Category A rotorcraft, 
it must be shown that a fan blade failure will not prevent continued 
safe flight either because of damage caused by the failed blade or loss 
of cooling air.
    (b) Category B. For cooling fans installed in category B rotorcraft, 
there

[[Page 797]]

must be means to protect the rotorcraft and allow a safe landing if a 
fan blade fails. It must be shown that--
    (1) The fan blade would be contained in the case of a failure;
    (2) Each fan is located so that a fan blade failure will not 
jeopardize safety; or
    (3) Each fan blade can withstand an ultimate load of 1.5 times the 
centrifugal force expected in service, limited by either--
    (i) The highest rotational speeds achievable under uncontrolled 
conditions; or
    (ii) An overspeed limiting device.
    (c) Fatigue evaluation. Unless a fatigue evaluation under Sec. 
29.571 is conducted, it must be shown that cooling fan blades are not 
operating at resonant conditions within the operating limits of the 
rotorcraft.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655 (c))

[Amdt. 29-13, 42 FR 15046, Mar. 17, 1977, as amended by Amdt. 29-26, 53 
FR 34215, Sept. 2, 1988]

                           Rotor Drive System



Sec. 29.917  Design.

    (a) General. The rotor drive system includes any part necessary to 
transmit power from the engines to the rotor hubs. This includes gear 
boxes, shafting, universal joints, couplings, rotor brake assemblies, 
clutches, supporting bearings for shafting, any attendant accessory pads 
or drives, and any cooling fans that are a part of, attached to, or 
mounted on the rotor drive system.
    (b) Design assessment. A design assessment must be performed to 
ensure that the rotor drive system functions safely over the full range 
of conditions for which certification is sought. The design assessment 
must include a detailed failure analysis to identify all failures that 
will prevent continued safe flight or safe landing and must identify the 
means to minimize the likelihood of their occurrence.
    (c) Arrangement. Rotor drive systems must be arranged as follows:
    (1) Each rotor drive system of multiengine rotorcraft must be 
arranged so that each rotor necessary for operation and control will 
continue to be driven by the remaining engines if any engine fails.
    (2) For single-engine rotorcraft, each rotor drive system must be so 
arranged that each rotor necessary for control in autorotation will 
continue to be driven by the main rotors after disengagement of the 
engine from the main and auxiliary rotors.
    (3) Each rotor drive system must incorporate a unit for each engine 
to automatically disengage that engine from the main and auxiliary 
rotors if that engine fails.
    (4) If a torque limiting device is used in the rotor drive system, 
it must be located so as to allow continued control of the rotorcraft 
when the device is operating.
    (5) If the rotors must be phased for intermeshing, each system must 
provide constant and positive phase relationship under any operating 
condition.
    (6) If a rotor dephasing device is incorporated, there must be means 
to keep the rotors locked in proper phase before operation.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55472, Dec. 20, 1976; Amdt. 29-40, 61 FR 21908, May 10, 1996]



Sec. 29.921  Rotor brake.

    If there is a means to control the rotation of the rotor drive 
system independently of the engine, any limitations on the use of that 
means must be specified, and the control for that means must be guarded 
to prevent inadvertent operation.



Sec. 29.923  Rotor drive system and control mechanism tests.

    (a) Endurance tests, general. Each rotor drive system and rotor 
control mechanism must be tested, as prescribed in paragraphs (b) 
through (n) and (p) of this section, for at least 200 hours plus the 
time required to meet the requirements of paragraphs (b)(2), (b)(3), and 
(k) of this section. These tests must be conducted as follows:
    (1) Ten-hour test cycles must be used, except that the test cycle 
must be extended to include the OEI test of paragraphs (b)(2) and (k), 
of this section if OEI ratings are requested.

[[Page 798]]

    (2) The tests must be conducted on the rotorcraft.
    (3) The test torque and rotational speed must be--
    (i) Determined by the powerplant limitations; and
    (ii) Absorbed by the rotors to be approved for the rotorcraft.
    (b) Endurance tests; takeoff run. The takeoff run must be conducted 
as follows:
    (1) Except as prescribed in paragraphs (b)(2) and (b)(3) of this 
section, the takeoff torque run must consist of 1 hour of alternate runs 
of 5 minutes at takeoff torque and the maximum speed for use with 
takeoff torque, and 5 minutes at as low an engine idle speed as 
practicable. The engine must be declutched from the rotor drive system, 
and the rotor brake, if furnished and so intended, must be applied 
during the first minute of the idle run. During the remaining 4 minutes 
of the idle run, the clutch must be engaged so that the engine drives 
the rotors at the minimum practical r.p.m. The engine and the rotor 
drive system must be accelerated at the maximum rate. When declutching 
the engine, it must be decelerated rapidly enough to allow the operation 
of the overrunning clutch.
    (2) For helicopters for which the use of a 2\1/2\-minute OEI rating 
is requested, the takeoff run must be conducted as prescribed in 
paragraph (b)(1) of this section, except for the third and sixth runs 
for which the takeoff torque and the maximum speed for use with takeoff 
torque are prescribed in that paragraph. For these runs, the following 
apply:
    (i) Each run must consist of at least one period of 2\1/2\ minutes 
with takeoff torque and the maximum speed for use with takeoff torque on 
all engines.
    (ii) Each run must consist of at least one period, for each engine 
in sequence, during which that engine simulates a power failure and the 
remaining engines are run at the 2\1/2\-minute OEI torque and the 
maximum speed for use with 2\1/2\-minute OEI torque for 2\1/2\ minutes.
    (3) For multiengine, turbine-powered rotorcraft for which the use of 
30-second/2-minute OEI power is requested, the takeoff run must be 
conducted as prescribed in paragraph (b)(1) of this section except for 
the following:
    (i) Immediately following any one 5-minute power-on run required by 
paragraph (b)(1) of this section, simulate a failure for each power 
source in turn, and apply the maximum torque and the maximum speed for 
use with 30-second OEI power to the remaining affected drive system 
power inputs for not less than 30 seconds. Each application of 30-second 
OEI power must be followed by two applications of the maximum torque and 
the maximum speed for use with the 2 minute OEI power for not less than 
2 minutes each; the second application must follow a period at 
stabilized continuous or 30 minute OEI power (whichever is requested by 
the applicant). At least one run sequence must be conducted from a 
simulated ``flight idle'' condition. When conducted on a bench test, the 
test sequence must be conducted following stabilization at take-off 
power.
    (ii) For the purpose of this paragraph, an affected power input 
includes all parts of the rotor drive system which can be adversely 
affected by the application of higher or asymmetric torque and speed 
prescribed by the test.
    (iii) This test may be conducted on a representative bench test 
facility when engine limitations either preclude repeated use of this 
power or would result in premature engine removals during the test. The 
loads, the vibration frequency, and the methods of application to the 
affected rotor drive system components must be representative of 
rotorcraft conditions. Test components must be those used to show 
compliance with the remainder of this section.
    (c) Endurance tests; maximum continuous run. Three hours of 
continuous operation at maximum continuous torque and the maximum speed 
for use with maximum continuous torque must be conducted as follows:
    (1) The main rotor controls must be operated at a minimum of 15 
times each hour through the main rotor pitch positions of maximum 
vertical thrust, maximum forward thrust component, maximum aft thrust 
component, maximum left thrust component, and maximum right thrust 
component, except that the control movements need not

[[Page 799]]

produce loads or blade flapping motion exceeding the maximum loads of 
motions encountered in flight.
    (2) The directional controls must be operated at a minimum of 15 
times each hour through the control extremes of maximum right turning 
torque, neutral torque as required by the power applied to the main 
rotor, and maximum left turning torque.
    (3) Each maximum control position must be held for at least 10 
seconds, and the rate of change of control position must be at least as 
rapid as that for normal operation.
    (d) Endurance tests; 90 percent of maximum continuous run. One hour 
of continuous operation at 90 percent of maximum continuous torque and 
the maximum speed for use with 90 percent of maximum continuous torque 
must be conducted.
    (e) Endurance tests; 80 percent of maximum continuous run. One hour 
of continuous operation at 80 percent of maximum continuous torque and 
the minimum speed for use with 80 percent of maximum continuous torque 
must be conducted.
    (f) Endurance tests; 60 percent of maximum continuous run. Two hours 
or, for helicopters for which the use of either 30-minute OEI power or 
continuous OEI power is requested, 1 hour of continuous operation at 60 
percent of maximum continuous torque and the minimum speed for use with 
60 percent of maximum continuous torque must be conducted.
    (g) Endurance tests; engine malfunctioning run. It must be 
determined whether malfunctioning of components, such as the engine fuel 
or ignition systems, or whether unequal engine power can cause dynamic 
conditions detrimental to the drive system. If so, a suitable number of 
hours of operation must be accomplished under those conditions, 1 hour 
of which must be included in each cycle, and the remaining hours of 
which must be accomplished at the end of the 20 cycles. If no 
detrimental condition results, an additional hour of operation in 
compliance with paragraph (b) of this section must be conducted in 
accordance with the run schedule of paragraph (b)(1) of this section 
without consideration of paragraph (b)(2) of this section.
    (h) Endurance tests; overspeed run. One hour of continuous operation 
must be conducted at maximum continuous torque and the maximum power-on 
overspeed expected in service, assuming that speed and torque limiting 
devices, if any, function properly.
    (i) Endurance tests; rotor control positions. When the rotor 
controls are not being cycled during the tie-down tests, the rotor must 
be operated, using the procedures prescribed in paragraph (c) of this 
section, to produce each of the maximum thrust positions for the 
following percentages of test time (except that the control positions 
need not produce loads or blade flapping motion exceeding the maximum 
loads or motions encountered in flight):
    (1) For full vertical thrust, 20 percent.
    (2) For the forward thrust component, 50 percent.
    (3) For the right thrust component, 10 percent.
    (4) For the left thrust component, 10 percent.
    (5) For the aft thrust component, 10 percent.
    (j) Endurance tests, clutch and brake engagements. A total of at 
least 400 clutch and brake engagements, including the engagements of 
paragraph (b) of this section, must be made during the takeoff torque 
runs and, if necessary, at each change of torque and speed throughout 
the test. In each clutch engagement, the shaft on the driven side of the 
clutch must be accelerated from rest. The clutch engagements must be 
accomplished at the speed and by the method prescribed by the applicant. 
During deceleration after each clutch engagement, the engines must be 
stopped rapidly enough to allow the engines to be automatically 
disengaged from the rotors and rotor drives. If a rotor brake is 
installed for stopping the rotor, the clutch, during brake engagements, 
must be disengaged above 40 percent of maximum continuous rotor speed 
and the rotors allowed to decelerate to 40 percent of maximum continuous 
rotor speed, at which time the rotor brake must be applied. If the 
clutch design does not allow stopping the rotors with the engine 
running, or if no clutch is provided, the engine must be stopped

[[Page 800]]

before each application of the rotor brake, and then immediately be 
started after the rotors stop.
    (k) Endurance tests; OEI power run--(1) 30-minute OEI power run. For 
rotorcraft for which the use of 30-minute OEI power is requested, a run 
at 30-minute OEI torque and the maximum speed for use with 30-minute OEI 
torque must be conducted as follows: For each engine, in sequence, that 
engine must be inoperative and the remaining engines must be run for a 
30-minute period.
    (2) Continuous OEI power run. For rotorcraft for which the use of 
continuous OEI power is requested, a run at continuous OEI torque and 
the maximum speed for use with continuous OEI torque must be conducted 
as follows: For each engine, in sequence, that engine must be 
inoperative and the remaining engines must be run for 1 hour.
    (3) The number of periods prescribed in paragraph (k)(1) or (k)(2) 
of this section may not be less than the number of engines, nor may it 
be less than two.
    (l) [Reserved]
    (m) Any components that are affected by maneuvering and gust loads 
must be investigated for the same flight conditions as are the main 
rotors, and their service lives must be determined by fatigue tests or 
by other acceptable methods. In addition, a level of safety equal to 
that of the main rotors must be provided for--
    (1) Each component in the rotor drive system whose failure would 
cause an uncontrolled landing;
    (2) Each component essential to the phasing of rotors on multirotor 
rotorcraft, or that furnishes a driving link for the essential control 
of rotors in autorotation; and
    (3) Each component common to two or more engines on multiengine 
rotorcraft.
    (n) Special tests. Each rotor drive system designed to operate at 
two or more gear ratios must be subjected to special testing for 
durations necessary to substantiate the safety of the rotor drive 
system.
    (o) Each part tested as prescribed in this section must be in a 
serviceable condition at the end of the tests. No intervening 
disassembly which might affect test results may be conducted.
    (p) Endurance tests; operating lubricants. To be approved for use in 
rotor drive and control systems, lubricants must meet the specifications 
of lubricants used during the tests prescribed by this section. 
Additional or alternate lubricants may be qualified by equivalent 
testing or by comparative analysis of lubricant specifications and rotor 
drive and control system characteristics. In addition--
    (1) At least three 10-hour cycles required by this section must be 
conducted with transmission and gearbox lubricant temperatures, at the 
location prescribed for measurement, not lower than the maximum 
operating temperature for which approval is requested;
    (2) For pressure lubricated systems, at least three 10-hour cycles 
required by this section must be conducted with the lubricant pressure, 
at the location prescribed for measurement, not higher than the minimum 
operating pressure for which approval is requested; and
    (3) The test conditions of paragraphs (p)(1) and (p)(2) of this 
section must be applied simultaneously and must be extended to include 
operation at any one-engine-inoperative rating for which approval is 
requested.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-1, 30 
FR 8778, July 13, 1965; Amdt. 29-17, 43 FR 50600, Oct. 30, 1978; Amdt. 
29-26, 53 FR 34215, Sept. 2, 1988; Amdt. 29-31, 55 FR 38967, Sept. 21, 
1990; Amdt. 29-34, 59 FR 47768, Sept. 16, 1994; Amdt. 29-40, 61 FR 
21908, May 10, 1996; Amdt. 29-42, 63 FR 43285, Aug. 12, 1998]



Sec. 29.927  Additional tests.

    (a) Any additional dynamic, endurance, and operational tests, and 
vibratory investigations necessary to determine that the rotor drive 
mechanism is safe, must be performed.
    (b) If turbine engine torque output to the transmission can exceed 
the highest engine or transmission torque limit, and that output is not 
directly controlled by the pilot under normal operating conditions (such 
as where

[[Page 801]]

the primary engine power control is accomplished through the flight 
control), the following test must be made:
    (1) Under conditions associated with all engines operating, make 200 
applications, for 10 seconds each, of torque that is at least equal to 
the lesser of--
    (i) The maximum torque used in meeting Sec. 29.923 plus 10 percent; 
or
    (ii) The maximum torque attainable under probable operating 
conditions, assuming that torque limiting devices, if any, function 
properly.
    (2) For multiengine rotorcraft under conditions associated with each 
engine, in turn, becoming inoperative, apply to the remaining 
transmission torque inputs the maximum torque attainable under probable 
operating conditions, assuming that torque limiting devices, if any, 
function properly. Each transmission input must be tested at this 
maximum torque for at least fifteen minutes.
    (c) Lubrication system failure. For lubrication systems required for 
proper operation of rotor drive systems, the following apply:
    (1) Category A. Unless such failures are extremely remote, it must 
be shown by test that any failure which results in loss of lubricant in 
any normal use lubrication system will not prevent continued safe 
operation, although not necessarily without damage, at a torque and 
rotational speed prescribed by the applicant for continued flight, for 
at least 30 minutes after perception by the flightcrew of the 
lubrication system failure or loss of lubricant.
    (2) Category B. The requirements of Category A apply except that the 
rotor drive system need only be capable of operating under autorotative 
conditions for at least 15 minutes.
    (d) Overspeed test. The rotor drive system must be subjected to 50 
overspeed runs, each 30 3 seconds in duration, at 
not less than either the higher of the rotational speed to be expected 
from an engine control device failure or 105 percent of the maximum 
rotational speed, including transients, to be expected in service. If 
speed and torque limiting devices are installed, are independent of the 
normal engine control, and are shown to be reliable, their rotational 
speed limits need not be exceeded. These runs must be conducted as 
follows:
    (1) Overspeed runs must be alternated with stabilizing runs of from 
1 to 5 minutes duration each at 60 to 80 percent of maximum continuous 
speed.
    (2) Acceleration and deceleration must be accomplished in a period 
not longer than 10 seconds (except where maximum engine acceleration 
rate will require more than 10 seconds), and the time for changing 
speeds may not be deducted from the specified time for the overspeed 
runs.
    (3) Overspeed runs must be made with the rotors in the flattest 
pitch for smooth operation.
    (e) The tests prescribed in paragraphs (b) and (d) of this section 
must be conducted on the rotorcraft and the torque must be absorbed by 
the rotors to be installed, except that other ground or flight test 
facilities with other appropriate methods of torque absorption may be 
used if the conditions of support and vibration closely simulate the 
conditions that would exist during a test on the rotorcraft.
    (f) Each test prescribed by this section must be conducted without 
intervening disassembly and, except for the lubrication system failure 
test required by paragraph (c) of this section, each part tested must be 
in a serviceable condition at the conclusion of the test.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Amdt. 29-3, 33 FR 969, Jan. 26, 1968, as amended by Amdt. 29-17, 43 FR 
50601, Oct. 30, 1978; Amdt. 29-26, 53 FR 34216, Sept. 2, 1988]



Sec. 29.931  Shafting critical speed.

    (a) The critical speeds of any shafting must be determined by 
demonstration except that analytical methods may be used if reliable 
methods of analysis are available for the particular design.
    (b) If any critical speed lies within, or close to, the operating 
ranges for idling, power-on, and autorotative conditions, the stresses 
occurring at that speed must be within safe limits. This must be shown 
by tests.
    (c) If analytical methods are used and show that no critical speed 
lies within the permissible operating ranges, the

[[Page 802]]

margins between the calculated critical speeds and the limits of the 
allowable operating ranges must be adequate to allow for possible 
variations between the computed and actual values.

[Amdt. 29-12, 41 FR 55472, Dec. 20, 1976]



Sec. 29.935  Shafting joints.

    Each universal joint, slip joint, and other shafting joints whose 
lubrication is necessary for operation must have provision for 
lubrication.



Sec. 29.939  Turbine engine operating characteristics.

    (a) Turbine engine operating characteristics must be investigated in 
flight to determine that no adverse characteristics (such as stall, 
surge, of flameout) are present, to a hazardous degree, during normal 
and emergency operation within the range of operating limitations of the 
rotorcraft and of the engine.
    (b) The turbine engine air inlet system may not, as a result of 
airflow distortion during normal operation, cause vibration harmful to 
the engine.
    (c) For governor-controlled engines, it must be shown that there 
exists no hazardous torsional instability of the drive system associated 
with critical combinations of power, rotational speed, and control 
displacement.

[Amdt. 29-2, 32 FR 6914, May 5, 1967, as amended by Amdt. 29-12, 41 FR 
55473, Dec. 20, 1976]

                               Fuel System



Sec. 29.951  General.

    (a) Each fuel system must be constructed and arranged to ensure a 
flow of fuel at a rate and pressure established for proper engine and 
auxiliary power unit functioning under any likely operating conditions, 
including the maneuvers for which certification is requested and during 
which the engine or auxiliary power unit is permitted to be in 
operation.
    (b) Each fuel system must be arranged so that--
    (1) No engine or fuel pump can draw fuel from more than one tank at 
a time; or
    (2) There are means to prevent introducing air into the system.
    (c) Each fuel system for a turbine engine must be capable of 
sustained operation throughout its flow and pressure range with fuel 
initially saturated with water at 80 degrees F. and having 0.75cc of 
free water per gallon added and cooled to the most critical condition 
for icing likely to be encountered in operation.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-10, 39 
FR 35462, Oct. 1, 1974; Amdt. 29-12, 41 FR 55473, Dec. 20, 1976]



Sec. 29.952  Fuel system crash resistance.

    Unless other means acceptable to the Administrator are employed to 
minimize the hazard of fuel fires to occupants following an otherwise 
survivable impact (crash landing), the fuel systems must incorporate the 
design features of this section. These systems must be shown to be 
capable of sustaining the static and dynamic deceleration loads of this 
section, considered as ultimate loads acting alone, measured at the 
system component's center of gravity without structural damage to the 
system components, fuel tanks, or their attachments that would leak fuel 
to an ignition source.
    (a) Drop test requirements. Each tank, or the most critical tank, 
must be drop-tested as follows:
    (1) The drop height must be at least 50 feet.
    (2) The drop impact surface must be nondeforming.
    (3) The tanks must be filled with water to 80 percent of the normal, 
full capacity.
    (4) The tank must be enclosed in a surrounding structure 
representative of the installation unless it can be established that the 
surrounding structure is free of projections or other design features 
likely to contribute to upture of the tank.
    (5) The tank must drop freely and impact in a horizontal position 
10[deg].
    (6) After the drop test, there must be no leakage.
    (b) Fuel tank load factors. Except for fuel tanks located so that 
tank rupture with fuel release to either significant ignition sources, 
such as engines, heaters, and auxiliary power units, or occupants is 
extremely remote, each fuel tank must be designed and installed to 
retain its contents under the following

[[Page 803]]

ultimate inertial load factors, acting alone.
    (1) For fuel tanks in the cabin:
    (i) Upward--4g.
    (ii) Forward--16g.
    (iii) Sideward--8g.
    (iv) Downward--20g.
    (2) For fuel tanks located above or behind the crew or passenger 
compartment that, if loosened, could injure an occupant in an emergency 
landing:
    (i) Upward--1.5g.
    (ii) Forward--8g.
    (iii) Sideward--2g.
    (iv) Downward--4g.
    (3) For fuel tanks in other areas:
    (i) Upward--1.5g.
    (ii) Forward--4g.
    (iii) Sideward--2g.
    (iv) Downward--4g.
    (c) Fuel line self-sealing breakaway couplings. Self-sealing 
breakaway couplings must be installed unless hazardous relative motion 
of fuel system components to each other or to local rotorcraft structure 
is demonstrated to be extremely improbable or unless other means are 
provided. The couplings or equivalent devices must be installed at all 
fuel tank-to-fuel line connections, tank-to-tank interconnects, and at 
other points in the fuel system where local structural deformation could 
lead to the release of fuel.
    (1) The design and construction of self-sealing breakaway couplings 
must incorporate the following design features:
    (i) The load necessary to separate a breakaway coupling must be 
between 25 to 50 percent of the minimum ultimate failure load (ultimate 
strength) of the weakest component in the fluid-carrying line. The 
separation load must in no case be less than 300 pounds, regardless of 
the size of the fluid line.
    (ii) A breakaway coupling must separate whenever its ultimate load 
(as defined in paragraph (c)(1)(i) of this section) is applied in the 
failure modes most likely to occur.
    (iii) All breakaway couplings must incorporate design provisions to 
visually ascertain that the coupling is locked together (leak-free) and 
is open during normal installation and service.
    (iv) All breakaway couplings must incorporate design provisions to 
prevent uncoupling or unintended closing due to operational shocks, 
vibrations, or accelerations.
    (v) No breakaway coupling design may allow the release of fuel once 
the coupling has performed its intended function.
    (2) All individual breakaway couplings, coupling fuel feed systems, 
or equivalent means must be designed, tested, installed, and maintained 
so inadvertent fuel shutoff in flight is improbable in accordance with 
Sec. 29.955(a) and must comply with the fatigue evaluation requirements 
of Sec. 29.571 without leaking.
    (3) Alternate, equivalent means to the use of breakaway couplings 
must not create a survivable impact-induced load on the fuel line to 
which it is installed greater than 25 to 50 percent of the ultimate load 
(strength) of the weakest component in the line and must comply with the 
fatigue requirements of Sec. 29.571 without leaking.
    (d) Frangible or deformable structural attachments. Unless hazardous 
relative motion of fuel tanks and fuel system components to local 
rotorcraft structure is demonstrated to be extremely improbable in an 
otherwise survivable impact, frangible or locally deformable attachments 
of fuel tanks and fuel system components to local rotorcraft structure 
must be used. The attachment of fuel tanks and fuel system components to 
local rotorcraft structure, whether frangible or locally deformable, 
must be designed such that its separation or relative local deformation 
will occur without rupture or local tear-out of the fuel tank or fuel 
system component that will cause fuel leakage. The ultimate strength of 
frangible or deformable attachments must be as follows:
    (1) The load required to separate a frangible attachment from its 
support structure, or deform a locally deformable attachment relative to 
its support structure, must be between 25 and 50 percent of the minimum 
ultimate load (ultimate strength) of the weakest component in the 
attached system. In no case may the load be less than 300 pounds.
    (2) A frangible or locally deformable attachment must separate or 
locally deform as intended whenever its ultimate load (as defined in 
paragraph

[[Page 804]]

(d)(1) of this section) is applied in the modes most likely to occur.
    (3) All frangible or locally deformable attachments must comply with 
the fatigue requirements of Sec. 29.571.
    (e) Separation of fuel and ignition sources. To provide maximum 
crash resistance, fuel must be located as far as practicable from all 
occupiable areas and from all potential ignition sources.
    (f) Other basic mechanical design criteria. Fuel tanks, fuel lines, 
electrical wires, and electrical devices must be designed, constructed, 
and installed, as far as practicable, to be crash resistant.
    (g) Rigid or semirigid fuel tanks. Rigid or semirigid fuel tank or 
bladder walls must be impact and tear resistant.

[Doc. No. 26352, 59 FR 50387, Oct. 3, 1994]



Sec. 29.953  Fuel system independence.

    (a) For category A rotorcraft--
    (1) The fuel system must meet the requirements of Sec. 29.903(b); 
and
    (2) Unless other provisions are made to meet paragraph (a)(1) of 
this section, the fuel system must allow fuel to be supplied to each 
engine through a system independent of those parts of each system 
supplying fuel to other engines.
    (b) Each fuel system for a multiengine category B rotorcraft must 
meet the requirements of paragraph (a)(2) of this section. However, 
separate fuel tanks need not be provided for each engine.



Sec. 29.954  Fuel system lightning protection.

    The fuel system must be designed and arranged to prevent the 
ignition of fuel vapor within the system by--
    (a) Direct lightning strikes to areas having a high probability of 
stroke attachment;
    (b) Swept lightning strokes to areas where swept strokes are highly 
probable; and
    (c) Corona and streamering at fuel vent outlets.

[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]



Sec. 29.955  Fuel flow.

    (a) General. The fuel system for each engine must provide the engine 
with at least 100 percent of the fuel required under all operating and 
maneuvering conditions to be approved for the rotorcraft, including, as 
applicable, the fuel required to operate the engines under the test 
conditions required by Sec. 29.927. Unless equivalent methods are used, 
compliance must be shown by test during which the following provisions 
are met, except that combinations of conditions which are shown to be 
improbable need not be considered.
    (1) The fuel pressure, corrected for accelerations (load factors), 
must be within the limits specified by the engine type certificate data 
sheet.
    (2) The fuel level in the tank may not exceed that established as 
the unusable fuel supply for that tank under Sec. 29.959, plus that 
necessary to conduct the test.
    (3) The fuel head between the tank and the engine must be critical 
with respect to rotorcraft flight attitudes.
    (4) The fuel flow transmitter, if installed, and the critical fuel 
pump (for pump-fed systems) must be installed to produce (by actual or 
simulated failure) the critical restriction to fuel flow to be expected 
from component failure.
    (5) Critical values of engine rotational speed, electrical power, or 
other sources of fuel pump motive power must be applied.
    (6) Critical values of fuel properties which adversely affect fuel 
flow are applied during demonstrations of fuel flow capability.
    (7) The fuel filter required by Sec. 29.997 is blocked to the 
degree necessary to simulate the accumulation of fuel contamination 
required to activate the indicator required by Sec. 29.1305(a)(17).
    (b) Fuel transfer system. If normal operation of the fuel system 
requires fuel to be transferred to another tank, the transfer must occur 
automatically via a system which has been shown to maintain the fuel 
level in the receiving tank within acceptable limits during flight or 
surface operation of the rotorcraft.
    (c) Multiple fuel tanks. If an engine can be supplied with fuel from 
more than one tank, the fuel system, in addition to having appropriate 
manual switching capability, must be designed to prevent interruption of 
fuel flow to that engine, without attention by the flightcrew, when any 
tank supplying fuel to that engine is depleted of usable

[[Page 805]]

fuel during normal operation and any other tank that normally supplies 
fuel to that engine alone contains usable fuel.

[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]



Sec. 29.957  Flow between interconnected tanks.

    (a) Where tank outlets are interconnected and allow fuel to flow 
between them due to gravity or flight accelerations, it must be 
impossible for fuel to flow between tanks in quantities great enough to 
cause overflow from the tank vent in any sustained flight condition.
    (b) If fuel can be pumped from one tank to another in flight--
    (1) The design of the vents and the fuel transfer system must 
prevent structural damage to tanks from overfilling; and
    (2) There must be means to warn the crew before overflow through the 
vents occurs.



Sec. 29.959  Unusable fuel supply.

    The unusable fuel supply for each tank must be established as not 
less than the quantity at which the first evidence of malfunction occurs 
under the most adverse fuel feed condition occurring under any intended 
operations and flight maneuvers involving that tank.



Sec. 29.961  Fuel system hot weather operation.

    Each suction lift fuel system and other fuel systems conducive to 
vapor formation must be shown to operate satisfactorily (within 
certification limits) when using fuel at the most critical temperature 
for vapor formation under critical operating conditions including, if 
applicable, the engine operating conditions defined by Sec. 
29.927(b)(1) and (b)(2).

[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]



Sec. 29.963  Fuel tanks: general.

    (a) Each fuel tank must be able to withstand, without failure, the 
vibration, inertia, fluid, and structural loads to which it may be 
subjected in operation.
    (b) Each flexible fuel tank bladder or liner must be approved or 
shown to be suitable for the particular application and must be puncture 
resistant. Puncture resistance must be shown by meeting the TSO-C80, 
paragraph 16.0, requirements using a minimum puncture force of 370 
pounds.
    (c) Each integral fuel tank must have facilities for inspection and 
repair of its interior.
    (d) The maximum exposed surface temperature of all components in the 
fuel tank must be less by a safe margin than the lowest expected 
autoignition temperature of the fuel or fuel vapor in the tank. 
Compliance with this requirement must be shown under all operating 
conditions and under all normal or malfunction conditions of all 
components inside the tank.
    (e) Each fuel tank installed in personnel compartments must be 
isolated by fume-proof and fuel-proof enclosures that are drained and 
vented to the exterior of the rotorcraft. The design and construction of 
the enclosures must provide necessary protection for the tank, must be 
crash resistant during a survivable impact in accordance with Sec. 
29.952, and must be adequate to withstand loads and abrasions to be 
expected in personnel compartments.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 
FR 34217, Sept. 2, 1988; Amdt. 29-35, 59 FR 50388, Oct. 3, 1994]



Sec. 29.965  Fuel tank tests.

    (a) Each fuel tank must be able to withstand the applicable pressure 
tests in this section without failure or leakage. If practicable, test 
pressures may be applied in a manner simulating the pressure 
distribution in service.
    (b) Each conventional metal tank, each nonmetallic tank with walls 
that are not supported by the rotorcraft structure, and each integral 
tank must be subjected to a pressure of 3.5 p.s.i. unless the pressure 
developed during maximum limit acceleration or emergency deceleration 
with a full tank exceeds this value, in which case a hydrostatic head, 
or equivalent test, must be applied to duplicate the acceleration loads 
as far as possible. However, the pressure need not exceed 3.5 p.s.i. on 
surfaces not exposed to the acceleration loading.
    (c) Each nonmetallic tank with walls supported by the rotorcraft 
structure

[[Page 806]]

must be subjected to the following tests:
    (1) A pressure test of at least 2.0 p.s.i. This test may be 
conducted on the tank alone in conjunction with the test specified in 
paragraph (c)(2) of this section.
    (2) A pressure test, with the tank mounted in the rotorcraft 
structure, equal to the load developed by the reaction of the contents, 
with the tank full, during maximum limit acceleration or emergency 
deceleration. However, the pressure need not exceed 2.0 p.s.i. on 
surfaces faces not exposed to the acceleration loading.
    (d) Each tank with large unsupported or unstiffened flat areas, or 
with other features whose failure or deformation could cause leakage, 
must be subjected to the following test or its equivalent:
    (1) Each complete tank assembly and its supports must be vibration 
tested while mounted to simulate the actual installation.
    (2) The tank assembly must be vibrated for 25 hours while two-thirds 
full of any suitable fluid. The amplitude of vibration may not be less 
than one thirty-second of an inch, unless otherwise substantiated.
    (3) The test frequency of vibration must be as follows:
    (i) If no frequency of vibration resulting from any r.p.m. within 
the normal operating range of engine or rotor system speeds is critical, 
the test frequency of vibration, in number of cycles per minute, must, 
unless a frequency based on a more rational analysis is used, be the 
number obtained by averaging the maximum and minimum power-on engine 
speeds (r.p.m.) for reciprocating engine powered rotorcraft or 2,000 
c.p.m. for turbine engine powered rotorcraft.
    (ii) If only one frequency of vibration resulting from any r.p.m. 
within the normal operating range of engine or rotor system speeds is 
critical, that frequency of vibration must be the test frequency.
    (iii) If more than one frequency of vibration resulting from any 
r.p.m. within the normal operating range of engine or rotor system 
speeds is critical, the most critical of these frequencies must be the 
test frequency.
    (4) Under paragraph (d)(3)(ii) and (iii), the time of test must be 
adjusted to accomplish the same number of vibration cycles as would be 
accomplished in 25 hours at the frequency specified in paragraph 
(d)(3)(i) of this section.
    (5) During the test, the tank assembly must be rocked at the rate of 
16 to 20 complete cycles per minute through an angle of 15 degrees on 
both sides of the horizontal (30 degrees total), about the most critical 
axis, for 25 hours. If motion about more than one axis is likely to be 
critical, the tank must be rocked about each critical axis for 12\1/2\ 
hours.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655 (c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 
FR 15046, Mar. 17, 1977]



Sec. 29.967  Fuel tank installation.

    (a) Each fuel tank must be supported so that tank loads are not 
concentrated on unsupported tank surfaces. In addition--
    (1) There must be pads, if necessary, to prevent chafing between 
each tank and its supports;
    (2) The padding must be nonabsorbent or treated to prevent the 
absorption of fuel;
    (3) If flexible tank liners are used, they must be supported so that 
they are not required to withstand fluid loads; and
    (4) Each interior surface of tank compartments must be smooth and 
free of projections that could cause wear of the liner, unless--
    (i) There are means for protection of the liner at those points; or
    (ii) The construction of the liner itself provides such protection.
    (b) Any spaces adjacent to tank surfaces must be adequately 
ventilated to avoid accumulation of fuel or fumes in those spaces due to 
minor leakage. If the tank is in a sealed compartment, ventilation may 
be limited to drain holes that prevent clogging and that prevent 
excessive pressure resulting from altitude changes. If flexible tank 
liners are installed, the venting arrangement for the spaces between the 
liner and its container must maintain the proper relationship to tank 
vent

[[Page 807]]

pressures for any expected flight condition.
    (c) The location of each tank must meet the requirements of Sec. 
29.1185(b) and (c).
    (d) No rotorcraft skin immediately adjacent to a major air outlet 
from the engine compartment may act as the wall of an integral tank.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 
FR 34217, Sept. 2, 1988; Amdt. 29-35, 59 FR 50388, Oct. 3, 1994]



Sec. 29.969  Fuel tank expansion space.

    Each fuel tank or each group of fuel tanks with interconnected vent 
systems must have an expansion space of not less than 2 percent of the 
combined tank capacity. It must be impossible to fill the fuel tank 
expansion space inadvertently with the rotorcraft in the normal ground 
attitude.

[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]



Sec. 29.971  Fuel tank sump.

    (a) Each fuel tank must have a sump with a capacity of not less than 
the greater of--
    (1) 0.10 per cent of the tank capacity; or
    (2) \1/16\ gallon.
    (b) The capacity prescribed in paragraph (a) of this section must be 
effective with the rotorcraft in any normal attitude, and must be 
located so that the sump contents cannot escape through the tank outlet 
opening.
    (c) Each fuel tank must allow drainage of hazardous quantities of 
water from each part of the tank to the sump with the rotorcraft in any 
ground attitude to be expected in service.
    (d) Each fuel tank sump must have a drain that allows complete 
drainage of the sump on the ground.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55473, Dec. 20, 1976; Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]



Sec. 29.973  Fuel tank filler connection.

    (a) Each fuel tank filler connection must prevent the entrance of 
fuel into any part of the rotorcraft other than the tank itself during 
normal operations and must be crash resistant during a survivable impact 
in accordance with Sec. 29.952(c). In addition--
    (1) Each filler must be marked as prescribed in Sec. 29.1557(c)(1);
    (2) Each recessed filler connection that can retain any appreciable 
quantity of fuel must have a drain that discharges clear of the entire 
rotorcraft; and
    (3) Each filler cap must provide a fuel-tight seal under the fluid 
pressure expected in normal operation and in a survivable impact.
    (b) Each filler cap or filler cap cover must warn when the cap is 
not fully locked or seated on the filler connection.

[Doc. No. 26352, 59 FR 50388, Oct. 3, 1994]



Sec. 29.975  Fuel tank vents and carburetor vapor vents.

    (a) Fuel tank vents. Each fuel tank must be vented from the top part 
of the expansion space so that venting is effective under normal flight 
conditions. In addition--
    (1) The vents must be arranged to avoid stoppage by dirt or ice 
formation;
    (2) The vent arrangement must prevent siphoning of fuel during 
normal operation;
    (3) The venting capacity and vent pressure levels must maintain 
acceptable differences of pressure between the interior and exterior of 
the tank, during--
    (i) Normal flight operation;
    (ii) Maximum rate of ascent and descent; and
    (iii) Refueling and defueling (where applicable);
    (4) Airspaces of tanks with interconnected outlets must be 
interconnected;
    (5) There may be no point in any vent line where moisture can 
accumulate with the rotorcraft in the ground attitude or the level 
flight attitude, unless drainage is provided;
    (6) No vent or drainage provision may end at any point--
    (i) Where the discharge of fuel from the vent outlet would 
constitute a fire hazard; or
    (ii) From which fumes could enter personnel compartments; and
    (7) The venting system must be designed to minimize spillage of fuel 
through the vents to an ignition source

[[Page 808]]

in the event of a rollover during landing, ground operations, or a 
survivable impact.
    (b) Carburetor vapor vents. Each carburetor with vapor elimination 
connections must have a vent line to lead vapors back to one of the fuel 
tanks. In addition--
    (1) Each vent system must have means to avoid stoppage by ice; and
    (2) If there is more than one fuel tank, and it is necessary to use 
the tanks in a definite sequence, each vapor vent return line must lead 
back to the fuel tank used for takeoff and landing.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 
FR 34217, Sept. 2, 1988; Amdt. 29-35, 59 FR 50388, Oct. 3, 1994; Amdt. 
29-42, 63 FR 43285, Aug. 12, 1998]



Sec. 29.977  Fuel tank outlet.

    (a) There must be a fuel strainer for the fuel tank outlet or for 
the booster pump. This strainer must--
    (1) For reciprocating engine powered airplanes, have 8 to 16 meshes 
per inch; and
    (2) For turbine engine powered airplanes, prevent the passage of any 
object that could restrict fuel flow or damage any fuel system 
component.
    (b) The clear area of each fuel tank outlet strainer must be at 
least five times the area of the outlet line.
    (c) The diameter of each strainer must be at least that of the fuel 
tank outlet.
    (d) Each finger strainer must be accessible for inspection and 
cleaning.

[Amdt. 29-12, 41 FR 55473, Dec. 20, 1976]



Sec. 29.979  Pressure refueling and fueling provisions below fuel level.

    (a) Each fueling connection below the fuel level in each tank must 
have means to prevent the escape of hazardous quantities of fuel from 
that tank in case of malfunction of the fuel entry valve.
    (b) For systems intended for pressure refueling, a means in addition 
to the normal means for limiting the tank content must be installed to 
prevent damage to the tank in case of failure of the normal means.
    (c) The rotorcraft pressure fueling system (not fuel tanks and fuel 
tank vents) must withstand an ultimate load that is 2.0 times the load 
arising from the maximum pressure, including surge, that is likely to 
occur during fueling. The maximum surge pressure must be established 
with any combination of tank valves being either intentionally or 
inadvertently closed.
    (d) The rotorcraft defueling system (not including fuel tanks and 
fuel tank vents) must withstand an ultimate load that is 2.0 times the 
load arising from the maximum permissible defueling pressure (positive 
or negative) at the rotorcraft fueling connection.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55473, Dec. 20, 1976]

                         Fuel System Components



Sec. 29.991  Fuel pumps.

    (a) Compliance with Sec. 29.955 must not be jeopardized by failure 
of--
    (1) Any one pump except pumps that are approved and installed as 
parts of a type certificated engine; or
    (2) Any component required for pump operation except the engine 
served by that pump.
    (b) The following fuel pump installation requirements apply:
    (1) When necessary to maintain the proper fuel pressure--
    (i) A connection must be provided to transmit the carburetor air 
intake static pressure to the proper fuel pump relief valve connection; 
and
    (ii) The gauge balance lines must be independently connected to the 
carburetor inlet pressure to avoid incorrect fuel pressure readings.
    (2) The installation of fuel pumps having seals or diaphragms that 
may leak must have means for draining leaking fuel.
    (3) Each drain line must discharge where it will not create a fire 
hazard.

[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]



Sec. 29.993  Fuel system lines and fittings.

    (a) Each fuel line must be installed and supported to prevent 
excessive vibration and to withstand loads due to fuel pressure, valve 
actuation, and accelerated flight conditions.
    (b) Each fuel line connected to components of the rotorcraft between

[[Page 809]]

which relative motion could exist must have provisions for flexibility.
    (c) Each flexible connection in fuel lines that may be under 
pressure or subjected to axial loading must use flexible hose 
assemblies.
    (d) Flexible hose must be approved.
    (e) No flexible hose that might be adversely affected by high 
temperatures may be used where excessive temperatures will exist during 
operation or after engine shutdown.



Sec. 29.995  Fuel valves.

    In addition to meeting the requirements of Sec. 29.1189, each fuel 
valve must--
    (a) [Reserved]
    (b) Be supported so that no loads resulting from their operation or 
from accelerated flight conditions are transmitted to the lines attached 
to the valve.

(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655 (c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 
FR 15046, Mar. 17, 1977]



Sec. 29.997  Fuel strainer or filter.

    There must be a fuel strainer or filter between the fuel tank outlet 
and the inlet of the first fuel system component which is susceptible to 
fuel contamination, including but not limited to the fuel metering 
device or an engine positive displacement pump, whichever is nearer the 
fuel tank outlet. This fuel strainer or filter must--
    (a) Be accessible for draining and cleaning and must incorporate a 
screen or element which is easily removable;
    (b) Have a sediment trap and drain, except that it need not have a 
drain if the strainer or filter is easily removable for drain purposes;
    (c) Be mounted so that its weight is not supported by the connecting 
lines or by the inlet or outlet connections of the strainer or filter 
inself, unless adequate strengh margins under all loading conditions are 
provided in the lines and connections; and
    (d) Provide a means to remove from the fuel any contaminant which 
would jeopardize the flow of fuel through rotorcraft or engine fuel 
system components required for proper rotorcraft or engine fuel system 
operation.

[Amdt. 29-10, 39 FR 35462, Oct. 1, 1974, as amended by Amdt. 29-22, 49 
FR 6850, Feb. 23, 1984; Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]



Sec. 29.999  Fuel system drains.

    (a) There must be at least one accessible drain at the lowest point 
in each fuel system to completely drain the system with the rotorcraft 
in any ground attitude to be expected in service.
    (b) Each drain required by paragraph (a) of this section including 
the drains prescribed in Sec. 29.971 must--
    (1) Discharge clear of all parts of the rotorcraft;
    (2) Have manual or automatic means to ensure positive closure in the 
off position; and
    (3) Have a drain valve--
    (i) That is readily accessible and which can be easily opened and 
closed; and
    (ii) That is either located or protected to prevent fuel spillage in 
the event of a landing with landing gear retracted.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55473, Dec. 20, 1976; Amdt. 29-26, 53 FR 34218, Sept. 2, 1988]



Sec. 29.1001  Fuel jettisoning.

    If a fuel jettisoning system is installed, the following apply:
    (a) Fuel jettisoning must be safe during all flight regimes for 
which jettisoning is to be authorized.
    (b) In showing compliance with paragraph (a) of this section, it 
must be shown that--
    (1) The fuel jettisoning system and its operation are free from fire 
hazard;
    (2) No hazard results from fuel or fuel vapors which impinge on any 
part of the rotorcraft during fuel jettisoning; and
    (3) Controllability of the rotorcraft remains satisfactory 
throughout the fuel jettisoning operation.
    (c) Means must be provided to automatically prevent jettisoning fuel 
below the level required for an all-engine climb at maximum continuous 
power from sea level to 5,000 feet altitude and cruise thereafter for 30 
minutes at maximum range engine power.

[[Page 810]]

    (d) The controls for any fuel jettisoning system must be designed to 
allow flight personnel (minimum crew) to safely interrupt fuel 
jettisoning during any part of the jettisoning operation.
    (e) The fuel jettisoning system must be designed to comply with the 
powerplant installation requirements of Sec. 29.901(c).
    (f) An auxiliary fuel jettisoning system which meets the 
requirements of paragraphs (a), (b), (d), and (e) of this section may be 
installed to jettison additional fuel provided it has separate and 
independent controls.

[Amdt. 29-26, 53 FR 34218, Sept. 2, 1988]

                               Oil System



Sec. 29.1011  Engines: general.

    (a) Each engine must have an independent oil system that can supply 
it with an appropriate quantity of oil at a temperature not above that 
safe for continuous operation.
    (b) The usable oil capacity of each system may not be less than the 
product of the endurance of the rotorcraft under critical operating 
conditions and the maximum allowable oil consumption of the engine under 
the same conditions, plus a suitable margin to ensure adequate 
circulation and cooling. Instead of a rational analysis of endurance and 
consumption, a usable oil capacity of one gallon for each 40 gallons of 
usable fuel may be used for reciprocating engine installations.
    (c) Oil-fuel ratios lower than those prescribed in paragraph (c) of 
this section may be used if they are substantiated by data on the oil 
consumption of the engine.
    (d) The ability of the engine and oil cooling provisions to maintain 
the oil temperature at or below the maximum established value must be 
shown under the applicable requirements of Sec. Sec. 29.1041 through 
29.1049.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 
FR 34218, Sept. 2, 1988]



Sec. 29.1013  Oil tanks.

    (a) Installation. Each oil tank installation must meet the 
requirements of Sec. 29.967.
    (b) Expansion space. Oil tank expansion space must be provided so 
that--
    (1) Each oil tank used with a reciprocating engine has an expansion 
space of not less than the greater of 10 percent of the tank capacity or 
0.5 gallon, and each oil tank used with a turbine engine has an 
expansion space of not less than 10 percent of the tank capacity;
    (2) Each reserve oil tank not directly connected to any engine has 
an expansion space of not less than two percent of the tank capacity; 
and
    (3) It is impossible to fill the expansion space inadvertently with 
the rotorcraft in the normal ground attitude.
    (c) Filler connections. Each recessed oil tank filler connection 
that can retain any appreciable quantity of oil must have a drain that 
discharges clear of the entire rotorcraft. In addition--
    (1) Each oil tank filler cap must provide an oil-tight seal under 
the pressure expected in operation;
    (2) For category A rotorcraft, each oil tank filler cap or filler 
cap cover must incorporate features that provide a warning when caps are 
not fully locked or seated on the filler connection; and
    (3) Each oil filler must be marked under Sec. 29.1557(c)(2).
    (d) Vent. Oil tanks must be vented as follows:
    (1) Each oil tank must be vented from the top part of the expansion 
space to that venting is effective under all normal flight conditions.
    (2) Oil tank vents must be arranged so that condensed water vapor 
that might freeze and obstruct the line cannot accumulate at any point;
    (e) Outlet. There must be means to prevent entrance into the tank 
itself, or into the tank outlet, of any object that might obstruct the 
flow of oil through the system. No oil tank outlet may be enclosed by a 
screen or guard that would reduce the flow of oil below a safe value at 
any operating temperature. There must be a shutoff valve at the outlet 
of each oil tank used with a turbine engine unless the external portion 
of the oil system (including oil tank supports) is fireproof.
    (f) Flexible liners. Each flexible oil tank liner must be approved 
or shown

[[Page 811]]

to be suitable for the particular installation.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-10, 39 
FR 35462, Oct. 1, 1974]



Sec. 29.1015  Oil tank tests.

    Each oil tank must be designed and installed so that--
    (a) It can withstand, without failure, any vibration, inertia, and 
fluid loads to which it may be subjected in operation; and
    (b) It meets the requirements of Sec. 29.965, except that instead 
of the pressure specified in Sec. 29.965(b)--
    (1) For pressurized tanks used with a turbine engine, the test 
pressure may not be less than 5 p.s.i. plus the maximum operating 
pressure of the tank; and
    (2) For all other tanks, the test pressure may not be less than 5 
p.s.i.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-10, 39 
FR 35462, Oct. 1, 1974]



Sec. 29.1017  Oil lines and fittings.

    (a) Each oil line must meet the requirements of Sec. 29.993.
    (b) Breather lines must be arranged so that--
    (1) Condensed water vapor that might freeze and obstruct the line 
cannot accumulate at any point;
    (2) The breather discharge will not constitute a fire hazard if 
foaming occurs, or cause emitted oil to strike the pilot's windshield; 
and
    (3) The breather does not discharge into the engine air induction 
system.



Sec. 29.1019  Oil strainer or filter.

    (a) Each turbine engine installation must incorporate an oil 
strainer or filter through which all of the engine oil flows and which 
meets the following requirements:
    (1) Each oil strainer or filter that has a bypass must be 
constructed and installed so that oil will flow at the normal rate 
through the rest of the system with the strainer or filter completely 
blocked.
    (2) The oil strainer or filter must have the capacity (with respect 
to operating limitations established for the engine) to ensure that 
engine oil system functioning is not impaired when the oil is 
contaminated to a degree (with respect to particle size and density) 
that is greater than that established for the engine under Part 33 of 
this chapter.
    (3) The oil strainer or filter, unless it is installed at an oil 
tank outlet, must incorporate a means to indicate contamination before 
it reaches the capacity established in accordance with paragraph (a)(2) 
of this section.
    (4) The bypass of a strainer or filter must be constructed and 
installed so that the release of collected contaminants is minimized by 
appropriate location of the bypass to ensure that collected contaminants 
are not in the bypass flow path.
    (5) An oil strainer or filter that has no bypass, except one that is 
installed at an oil tank outlet, must have a means to connect it to the 
warning system required in Sec. 29.1305(a)(18).
    (b) Each oil strainer or filter in a powerplant installation using 
reciprocating engines must be constructed and installed so that oil will 
flow at the normal rate through the rest of the system with the strainer 
or filter element completely blocked.

[Amdt. 29-10, 39 FR 35463, Oct. 1, 1974, as amended by Amdt. 29-22, 49 
FR 6850, Feb. 23, 1984; Amdt. 29-26, 53 FR 34218, Sept. 2, 1988]



Sec. 29.1021  Oil system drains.

    A drain (or drains) must be provided to allow safe drainage of the 
oil system. Each drain must--
    (a) Be accessible; and
    (b) Have manual or automatic means for positive locking in the 
closed position.

[Amdt. 29-22, 49 FR 6850, Feb. 23, 1984]



Sec. 29.1023  Oil radiators.

    (a) Each oil radiator must be able to withstand any vibration, 
inertia, and oil pressure loads to which it would be subjected in 
operation.
    (b) Each oil radiator air duct must be located, or equipped, so 
that, in case of fire, and with the airflow as it would be with and 
without the engine operating, flames cannot directly strike the 
radiator.

[[Page 812]]



Sec. 29.1025  Oil valves.

    (a) Each oil shutoff must meet the requirements of Sec. 29.1189.
    (b) The closing of oil shutoffs may not prevent autorotation.
    (c) Each oil valve must have positive stops or suitable index 
provisions in the ``on'' and ``off'' positions and must be supported so 
that no loads resulting from its operation or from accelerated flight 
conditions are transmitted to the lines attached to the valve.



Sec. 29.1027  Transmission and gearboxes: general.

    (a) The oil system for components of the rotor drive system that 
require continuous lubrication must be sufficiently independent of the 
lubrication systems of the engine(s) to ensure--
    (1) Operation with any engine inoperative; and
    (2) Safe autorotation.
    (b) Pressure lubrication systems for transmissions and gearboxes 
must comply with the requirements of Sec. Sec. 29.1013, paragraphs (c), 
(d), and (f) only, 29.1015, 29.1017, 29.1021, 29.1023, and 29.1337(d). 
In addition, the system must have--
    (1) An oil strainer or filter through which all the lubricant flows, 
and must--
    (i) Be designed to remove from the lubricant any contaminant which 
may damage transmission and drive system components or impede the flow 
of lubricant to a hazardous degree; and
    (ii) Be equipped with a bypass constructed and installed so that--
    (A) The lubricant will flow at the normal rate through the rest of 
the system with the strainer or filter completely blocked; and
    (B) The release of collected contaminants is minimized by 
appropriate location of the bypass to ensure that collected contaminants 
are not in the bypass flowpath;
    (iii) Be equipped with a means to indicate collection of 
contaminants on the filter or strainer at or before opening of the 
bypass;
    (2) For each lubricant tank or sump outlet supplying lubrication to 
rotor drive systems and rotor drive system components, a screen to 
prevent entrance into the lubrication system of any object that might 
obstruct the flow of lubricant from the outlet to the filter required by 
paragraph (b)(1) of this section. The requirements of paragraph (b)(1) 
of this section do not apply to screens installed at lubricant tank or 
sump outlets.
    (c) Splash type lubrication systems for rotor drive system gearboxes 
must comply with Sec. Sec. 29.1021 and 29.1337(d).

[Amdt. 29-26, 53 FR 34218, Sept. 2, 1988]

                                 Cooling



Sec. 29.1041  General.

    (a) The powerplant and auxiliary power unit cooling provisions must 
be able to maintain the temperatures of powerplant components, engine 
fluids, and auxiliary power unit components and fluids within the 
temperature limits established for these components and fluids, under 
ground, water, and flight operating conditions for which certification 
is requested, and after normal engine or auxiliary power unit shutdown, 
or both.
    (b) There must be cooling provisions to maintain the fluid 
temperatures in any power transmission within safe values under any 
critical surface (ground or water) and flight operating conditions.
    (c) Except for ground-use-only auxiliary power units, compliance 
with paragraphs (a) and (b) of this section must be shown by flight 
tests in which the temperatures of selected powerplant component and 
auxiliary power unit component, engine, and transmission fluids are 
obtained under the conditions prescribed in those paragraphs.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 
FR 34218, Sept. 2, 1988]



Sec. 29.1043  Cooling tests.

    (a) General. For the tests prescribed in Sec. 29.1041(c), the 
following apply:
    (1) If the tests are conducted under conditions deviating from the 
maximum ambient atmospheric temperature specified in paragraph (b) of 
this section, the recorded powerplant temperatures must be corrected 
under paragraphs (c) and (d) of this section, unless a more rational 
correction method is applicable.

[[Page 813]]

    (2) No corrected temperature determined under paragraph (a)(1) of 
this section may exceed established limits.
    (3) The fuel used during the cooling tests must be of the minimum 
grade approved for the engines, and the mixture settings must be those 
used in normal operation.
    (4) The test procedures must be as prescribed in Sec. Sec. 29.1045 
through 29.1049.
    (5) For the purposes of the cooling tests, a temperature is 
``stabilized'' when its rate of change is less than 2 [deg]F per minute.
    (b) Maximum ambient atmospheric temperature. A maximum ambient 
atmospheric temperature corresponding to sea level conditions of at 
least 100 degrees F. must be established. The assumed temperature lapse 
rate is 3.6 degrees F. per thousand feet of altitude above sea level 
until a temperature of -69.7 degrees F. is reached, above which altitude 
the temperature is considered constant at -69.7 degrees F. However, for 
winterization installations, the applicant may select a maximum ambient 
atmospheric temperature corresponding to sea level conditions of less 
than 100 degrees F.
    (c) Correction factor (except cylinder barrels). Unless a more 
rational correction applies, temperatures of engine fluids and 
powerplant components (except cylinder barrels) for which temperature 
limits are established, must be corrected by adding to them the 
difference between the maximum ambient atmospheric temperature and the 
temperature of the ambient air at the time of the first occurrence of 
the maximum component or fluid temperature recorded during the cooling 
test.
    (d) Correction factor for cylinder barrel temperatures. Cylinder 
barrel temperatures must be corrected by adding to them 0.7 times the 
difference between the maximum ambient atmospheric temperature and the 
temperature of the ambient air at the time of the first occurrence of 
the maximum cylinder barrel temperature recorded during the cooling 
test.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of 
the Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55473, Dec. 20, 1976; Amdt. 29-15, 43 FR 2327, Jan. 16, 1978; Amdt. 
29-26, 53 FR 34218, Sept. 2, 1988]



Sec. 29.1045  Climb cooling test procedures.

    (a) Climb cooling tests must be conducted under this section for--
    (1) Category A rotorcraft; and
    (2) Multiengine category B rotorcraft for which certification is 
requested under the category A powerplant installation requirements, and 
under the requirements of Sec. 29.861(a) at the steady rate of climb or 
descent established under Sec. 29.67(b).
    (b) The climb or descent cooling tests must be conducted with the 
engine inoperative that produces the most adverse cooling conditions for 
the remaining engines and powerplant components.
    (c) Each operating engine must--
    (1) For helicopters for which the use of 30-minute OEI power is 
requested, be at 30-minute OEI power for 30 minutes, and then at maximum 
continuous power (or at full throttle when above the critical altitude);
    (2) For helicopters for which the use of continuous OEI power is 
requested, be at continuous OEI power (or at full throttle when above 
the critical altitude); and
    (3) For other rotorcraft, be at maximum continuous power (or at full 
throttle when above the critical altitude).
    (d) After temperatures have stabilized in flight, the climb must 
be--
    (1) Begun from an altitude not greater than the lower of--
    (i) 1,000 feet below the engine critcal altitude; and
    (ii) 1,000 feet below the maximum altitude at which the rate of 
climb is 150 f.p.m; and
    (2) Continued for at least five minutes after the occurrence of the 
highest temperature recorded, or until the

[[Page 814]]

rotorcraft reaches the maximum altitude for which certification is 
requested.
    (e) For category B rotorcraft without a positive rate of climb, the 
descent must begin at the all-engine-critical altitude and end at the 
higher of--
    (1) The maximum altitude at which level flight can be maintained 
with one engine operative; and
    (2) Sea level.
    (f) The climb or descent must be conducted at an airspeed 
representing a normal operational practice for the configuration being 
tested. However, if the cooling provisions are sensitive to rotorcraft 
speed, the most critical airspeed must be used, but need not exceed the 
speeds established under Sec. 29.67(a)(2) or Sec. 29.67(b). The climb 
cooling test may be conducted in conjunction with the takeoff cooling 
test of Sec. 29.1047.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 
FR 34218, Sept. 2, 1988]



Sec. 29.1047  Takeoff cooling test procedures.

    (a) Category A. For each category A rotorcraft, cooling must be 
shown during takeoff and subsequent climb as follows:
    (1) Each temperature must be stabilized while hovering in ground 
effect with--
    (i) The power necessary for hovering;
    (ii) The appropriate cowl flap and shutter settings; and
    (iii) The maximum weight.
    (2) After the temperatures have stabilized, a climb must be started 
at the lowest practicable altitude and must be conducted with one engine 
inoperative.
    (3) The operating engines must be at the greatest power for which 
approval is sought (or at full throttle when above the critical 
altitude) for the same period as this power is used in determining the 
takeoff climbout path under Sec. 29.59.
    (4) At the end of the time interval prescribed in paragraph (b)(3) 
of this section, the power must be changed to that used in meeting Sec. 
29.67(a)(2) and the climb must be continued for--
    (i) Thirty minutes, if 30-minute OEI power is used; or
    (ii) At least 5 minutes after the occurrence of the highest 
temperature recorded, if continuous OEI power or maximum continuous 
power is used.
    (5) The speeds must be those used in determining the takeoff flight 
path under Sec. 29.59.
    (b) Category B. For each category B rotorcraft, cooling must be 
shown during takeoff and subsequent climb as follows:
    (1) Each temperature must be stabilized while hovering in ground 
effect with--
    (i) The power necessary for hovering;
    (ii) The appropriate cowl flap and shutter settings; and
    (iii) The maximum weight.
    (2) After the temperatures have stabilized, a climb must be started 
at the lowest practicable altitude with takeoff power.
    (3) Takeoff power must be used for the same time interval as takeoff 
power is used in determining the takeoff flight path under Sec. 29.63.
    (4) At the end of the time interval prescribed in paragraph (a)(3) 
of this section, the power must be reduced to maximum continuous power 
and the climb must be continued for at least five minutes after the 
occurance of the highest temperature recorded.
    (5) The cooling test must be conducted at an airspeed corresponding 
to normal operating practice for the configuration being tested. 
However, if the cooling provisions are sensitive to rotorcraft speed, 
the most critical airspeed must be used, but need not exceed the speed 
for best rate of climb with maximum continuous power.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-1, 30 
FR 8778, July 13, 1965; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]



Sec. 29.1049  Hovering cooling test procedures.

    The hovering cooling provisions must be shown--
    (a) At maximum weight or at the greatest weight at which the 
rotorcraft can hover (if less), at sea level, with the power required to 
hover but not more than maximum continuous power, in the ground effect 
in still air, until at least five minutes after the occurrence of the 
highest temperature recorded; and

[[Page 815]]

    (b) With maximum continuous power, maximum weight, and at the 
altitude resulting in zero rate of climb for this configuration, until 
at least five minutes after the occurrence of the highest temperature 
recorded.

                            Induction System



Sec. 29.1091  Air induction.

    (a) The air induction system for each engine and auxiliary power 
unit must supply the air required by that engine and auxiliary power 
unit under the operating conditions for which certification is 
requested.
    (b) Each engine and auxiliary power unit air induction system must 
provide air for proper fuel metering and mixture distribution with the 
induction system valves in any position.
    (c) No air intake may open within the engine accessory section or 
within other areas of any powerplant compartment where emergence of 
backfire flame would constitute a fire hazard.
    (d) Each reciprocating engine must have an alternate air source.
    (e) Each alternate air intake must be located to prevent the 
entrance of rain, ice, or other foreign matter.
    (f) For turbine engine powered rotorcraft and rotorcraft 
incorporating auxiliary power units--
    (1) There must be means to prevent hazardous quantities of fuel 
leakage or overflow from drains, vents, or other components of flammable 
fluid systems from entering the engine or auxiliary power unit intake 
system; and
    (2) The air inlet ducts must be located or protected so as to 
minimize the ingestion of foreign matter during takeoff, landing, and 
taxiing.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 969, Jan. 26, 1968; Amdt. 29-17, 43 FR 50601, Oct. 30, 1978]



Sec. 29.1093  Induction system icing protection.

    (a) Reciprocating engines. Each reciprocating engine air induction 
system must have means to prevent and eliminate icing. Unless this is 
done by other means, it must be shown that, in air free of visible 
moisture at a temperature of 30 [deg]F., and with the engines at 60 
percent of maximum continuous power--
    (1) Each rotorcraft with sea level engines using conventional 
venturi carburetors has a preheater that can provide a heat rise of 90 
[deg]F.;
    (2) Each rotorcraft with sea level engines using carburetors tending 
to prevent icing has a preheater that can provide a heat rise of 70 
[deg]F.;
    (3) Each rotorcraft with altitude engines using conventional venturi 
carburetors has a preheater that can provide a heat rise of 120 [deg]F.; 
and
    (4) Each rotorcraft with altitude engines using carburetors tending 
to prevent icing has a preheater that can provide a heat rise of 100 
[deg]F.
    (b) Turbine engines. (1) It must be shown that each turbine engine 
and its air inlet system can operate throughout the flight power range 
of the engine (including idling)--
    (i) Without accumulating ice on engine or inlet system components 
that would adversely affect engine operation or cause a serious loss of 
power under the icing conditions specified in appendix C of this Part; 
and
    (ii) In snow, both falling and blowing, without adverse effect on 
engine operation, within the limitations established for the rotorcraft.
    (2) Each turbine engine must idle for 30 minutes on the ground, with 
the air bleed available for engine icing protection at its critical 
condition, without adverse effect, in an atmosphere that is at a 
temperature between 15[deg] and 30 [deg]F (between -9[deg] and -1 
[deg]C) and has a liquid water content not less than 0.3 grams per cubic 
meter in the form of drops having a mean effective diameter not less 
than 20 microns, followed by momentary operation at takeoff power or 
thrust. During the 30 minutes of idle operation, the engine may be run 
up periodically to a moderate power or thrust setting in a manner 
acceptable to the Administrator.
    (c) Supercharged reciprocating engines. For each engine having a 
supercharger to pressurize the air before it enters the carburetor, the 
heat rise in the air caused by that supercharging at any altitude may be 
utilized in determining compliance with paragraph (a) of this

[[Page 816]]

section if the heat rise utilized is that which will be available, 
automatically, for the applicable altitude and operation condition 
because of supercharging.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655 (c))

[Amdt. 29-3, 33 FR 969, Jan. 26, 1968, as amended by Amdt. 29-12, 41 FR 
55473, Dec. 20, 1976; Amdt. 29-13, 42 FR 15046, Mar. 17, 1977; Amdt. 29-
22, 49 FR 6850, Feb. 23, 1984; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]



Sec. 29.1101  Carburetor air preheater design.

    Each carburetor air preheater must be designed and constructed to--
    (a) Ensure ventilation of the preheater when the engine is operated 
in cold air;
    (b) Allow inspection of the exhaust manifold parts that it 
surrounds; and
    (c) Allow inspection of critical parts of the preheater itself.



Sec. 29.1103  Induction systems ducts and air duct systems.

    (a) Each induction system duct upstream of the first stage of the 
engine supercharger and of the auxiliary power unit compressor must have 
a drain to prevent the hazardous accumulation of fuel and moisture in 
the ground attitude. No drain may discharge where it might cause a fire 
hazard.
    (b) Each duct must be strong enough to prevent induction system 
failure from normal backfire conditions.
    (c) Each duct connected to components between which relative motion 
could exist must have means for flexibility.
    (d) Each duct within any fire zone for which a fire-extinguishing 
system is required must be at least--
    (1) Fireproof, if it passes through any firewall; or
    (2) Fire resistant, for other ducts, except that ducts for auxiliary 
power units must be fireproof within the auxiliary power unit fire zone.
    (e) Each auxiliary power unit induction system duct must be 
fireproof for a sufficient distance upstream of the auxiliary power unit 
compartment to prevent hot gas reverse flow from burning through 
auxiliary power unit ducts and entering any other compartment or area of 
the rotorcraft in which a hazard would be created resulting from the 
entry of hot gases. The materials used to form the remainder of the 
induction system duct and plenum chamber of the auxiliary power unit 
must be capable of resisting the maximum heat conditions likely to 
occur.
    (f) Each auxiliary power unit induction system duct must be 
constructed of materials that will not absorb or trap hazardous 
quantities of flammable fluids that could be ignited in the event of a 
surge or reverse flow condition.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-17, 43 
FR 50602, Oct. 30, 1978]



Sec. 29.1105  Induction system screens.

    If induction system screens are used--
    (a) Each screen must be upstream of the carburetor;
    (b) No screen may be in any part of the induction system that is the 
only passage through which air can reach the engine, unless it can be 
deiced by heated air;
    (c) No screen may be deiced by alcohol alone; and
    (d) It must be impossible for fuel to strike any screen.



Sec. 29.1107  Inter-coolers and after-coolers.

    Each inter-cooler and after-cooler must be able to withstand the 
vibration, inertia, and air pressure loads to which it would be 
subjected in operation.



Sec. 29.1109  Carburetor air cooling.

    It must be shown under Sec. 29.1043 that each installation using 
two-stage superchargers has means to maintain the air temperature, at 
the carburetor inlet, at or below the maximum established value.

[[Page 817]]

                             Exhaust System



Sec. 29.1121  General.

    For powerplant and auxiliary power unit installations the following 
apply:
    (a) Each exhaust system must ensure safe disposal of exhaust gases 
without fire hazard or carbon monoxide contamination in any personnel 
compartment.
    (b) Each exhaust system part with a surface hot enough to ignite 
flammable fluids or vapors must be located or shielded so that leakage 
from any system carrying flammable fluids or vapors will not result in a 
fire caused by impingement of the fluids or vapors on any part of the 
exhaust system including shields for the exhaust system.
    (c) Each component upon which hot exhaust gases could impinge, or 
that could be subjected to high temperatures from exhaust system parts, 
must be fireproof. Each exhaust system component must be separated by a 
fireproof shield from adjacent parts of the rotorcraft that are outside 
the engine and auxiliary power unit compartments.
    (d) No exhaust gases may discharge so as to cause a fire hazard with 
respect to any flammable fluid vent or drain.
    (e) No exhaust gases may discharge where they will cause a glare 
seriously affecting pilot vision at night.
    (f) Each exhaust system component must be ventilated to prevent 
points of excessively high temperature.
    (g) Each exhaust shroud must be ventilated or insulated to avoid, 
during normal operation, a temperature high enough to ignite any 
flammable fluids or vapors outside the shroud.
    (h) If significant traps exist, each turbine engine exhaust system 
must have drains discharging clear of the rotorcraft, in any normal 
ground and flight attitudes, to prevent fuel accumulation after the 
failure of an attempted engine start.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 755, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655 (c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 970, Jan. 26, 1968; Amdt. 29-13, 42 FR 15046, Mar. 17, 1977]



Sec. 29.1123  Exhaust piping.

    (a) Exhaust piping must be heat and corrosion resistant, and must 
have provisions to prevent failure due to expansion by operating 
temperatures.
    (b) Exhaust piping must be supported to withstand any vibration and 
inertia loads to which it would be subjected in operation.
    (c) Exhaust piping connected to components between which relative 
motion could exist must have provisions for flexibility.



Sec. 29.1125  Exhaust heat exchangers.

    For reciprocating engine powered rotorcraft the following apply:
    (a) Each exhaust heat exchanger must be constructed and installed to 
withstand the vibration, inertia, and other loads to which it would be 
subjected in operation. In addition--
    (1) Each exchanger must be suitable for continued operation at high 
temperatures and resistant to corrosion from exhaust gases;
    (2) There must be means for inspecting the critical parts of each 
exchanger;
    (3) Each exchanger must have cooling provisions wherever it is 
subject to contact with exhaust gases; and
    (4) No exhaust heat exchanger or muff may have stagnant areas or 
liquid traps that would increase the probability of ignition of 
flammable fluids or vapors that might be present in case of the failure 
or malfunction of components carrying flammable fluids.
    (b) If an exhaust heat exchanger is used for heating ventilating air 
used by personnel--
    (1) There must be a secondary heat exchanger between the primary 
exhaust gas heat exchanger and the ventilating air system; or
    (2) Other means must be used to prevent harmful contamination of the 
ventilating air.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55473, Dec. 20, 1976; Amdt. 29-41, 62 FR 46173, Aug. 29, 1997]

[[Page 818]]

                   Powerplant Controls and Accessories



Sec. 29.1141  Powerplant controls: general.

    (a) Powerplant controls must be located and arranged under Sec. 
29.777 and marked under Sec. 29.1555.
    (b) Each control must be located so that it cannot be inadvertently 
operated by persons entering, leaving, or moving normally in the 
cockpit.
    (c) Each flexible powerplant control must be approved.
    (d) Each control must be able to maintain any set position without--
    (1) Constant attention; or
    (2) Tendency to creep due to control loads or vibration.
    (e) Each control must be able to withstand operating loads without 
excessive deflection.
    (f) Controls of powerplant valves required for safety must have--
    (1) For manual valves, positive stops or in the case of fuel valves 
suitable index provisions, in the open and closed position; and
    (2) For power-assisted valves, a means to indicate to the flight 
crew when the valve--
    (i) Is in the fully open or fully closed position; or
    (ii) Is moving between the fully open and fully closed position.

(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 
FR 15046, Mar. 17, 1977; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]



Sec. 29.1142  Auxiliary power unit controls.

    Means must be provided on the flight deck for starting, stopping, 
and emergency shutdown of each installed auxiliary power unit.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Amdt. 29-17, 43 FR 50602, Oct. 30, 1978]



Sec. 29.1143  Engine controls.

    (a) There must be a separate power control for each engine.
    (b) Power controls must be arranged to allow ready synchronization 
of all engines by--
    (1) Separate control of each engine; and
    (2) Simultaneous control of all engines.
    (c) Each power control must provide a positive and immediately 
responsive means of controlling its engine.
    (d) Each fluid injection control other than fuel system control must 
be in the corresponding power control. However, the injection system 
pump may have a separate control.
    (e) If a power control incorporates a fuel shutoff feature, the 
control must have a means to prevent the inadvertent movement of the 
control into the shutoff position. The means must--
    (1) Have a positive lock or stop at the idle position; and
    (2) Require a separate and distinct operation to place the control 
in the shutoff position.
    (f) For rotorcraft to be certificated for a 30-second OEI power 
rating, a means must be provided to automatically activate and control 
the 30-second OEI power and prevent any engine from exceeding the 
installed engine limits associated with the 30-second OEI power rating 
approved for the rotorcraft.

[Amdt. 29-26, 53 FR 34219, Sept. 2, 1988, as amended by Amdt. 29-34, 59 
FR 47768, Sept. 16, 1994]



Sec. 29.1145  Ignition switches.

    (a) Ignition switches must control each ignition circuit on each 
engine.
    (b) There must be means to quickly shut off all ignition by the 
grouping of switches or by a master ignition control.
    (c) Each group of ignition switches, except ignition switches for 
turbine engines for which continuous ignition is not required, and each 
master ignition control must have a means to prevent its inadvertent 
operation.

(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655 (c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 
FR 15046, Mar. 17, 1977]

[[Page 819]]



Sec. 29.1147  Mixture controls.

    (a) If there are mixture controls, each engine must have a separate 
control, and the controls must be arranged to allow--
    (1) Separate control of each engine; and
    (2) Simultaneous control of all engines.
    (b) Each intermediate position of the mixture controls that 
corresponds to a normal operating setting must be identifiable by feel 
and sight.



Sec. 29.1151  Rotor brake controls.

    (a) It must be impossible to apply the rotor brake inadvertently in 
flight.
    (b) There must be means to warn the crew if the rotor brake has not 
been completely released before take off.



Sec. 29.1157  Carburetor air temperature controls.

    There must be a separate carburetor air temperature control for each 
engine.



Sec. 29.1159  Supercharger controls.

    Each supercharger control must be accessible to--
    (a) The pilots; or
    (b) (If there is a separate flight engineer station with a control 
panel) the flight engineer.



Sec. 29.1163  Powerplant accessories.

    (a) Each engine mounted accessory must--
    (1) Be approved for mounting on the engine involved;
    (2) Use the provisions on the engine for mounting; and
    (3) Be sealed in such a way as to prevent contamination of the 
engine oil system and the accessory system.
    (b) Electrical equipment subject to arcing or sparking must be 
installed, to minimize the probability of igniting flammable fluids or 
vapors.
    (c) If continued rotation of an engine-driven cabin supercharger or 
any remote accessory driven by the engine will be a hazard if they 
malfunction, there must be means to prevent their hazardous rotation 
without interfering with the continued operation of the engine.
    (d) Unless other means are provided, torque limiting means must be 
provided for accessory drives located on any component of the 
transmission and rotor drive system to prevent damage to these 
components from excessive accessory load.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-22, 49 
FR 6850, Feb. 23, 1984; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]



Sec. 29.1165  Engine ignition systems.

    (a) Each battery ignition system must be supplemented with a 
generator that is automatically available as an alternate source of 
electrical energy to allow continued engine operation if any battery 
becomes depleted.
    (b) The capacity of batteries and generators must be large enough to 
meet the simultaneous demands of the engine ignition system and the 
greatest demands of any electrical system components that draw from the 
same source.
    (c) The design of the engine ignition system must account for--
    (1) The condition of an inoperative generator;
    (2) The condition of a completely depleted battery with the 
generator running at its normal operating speed; and
    (3) The condition of a completely depleted battery with the 
generator operating at idling speed, if there is only one battery.
    (d) Magneto ground wiring (for separate ignition circuits) that lies 
on the engine side of any firewall must be installed, located, or 
protected, to minimize the probability of the simultaneous failure of 
two or more wires as a result of mechanical damage, electrical fault, or 
other cause.
    (e) No ground wire for any engine may be routed through a fire zone 
of another engine unless each part of that wire within that zone is 
fireproof.
    (f) Each ignition system must be independent of any electrical 
circuit that is not used for assisting, controlling, or analyzing the 
operation of that system.
    (g) There must be means to warn appropriate crewmembers if the 
malfunctioning of any part of the electrical

[[Page 820]]

system is causing the continuous discharge of any battery necessary for 
engine ignition.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55473, Dec. 20, 1976]

                       Powerplant Fire Protection



Sec. 29.1181  Designated fire zones: regions included.

    (a) Designated fire zones are--
    (1) The engine power section of reciprocating engines;
    (2) The engine accessory section of reciprocating engines;
    (3) Any complete powerplant compartment in which there is no 
isolation between the engine power section and the engine accessory 
section, for reciprocating engines;
    (4) Any auxiliary power unit compartment;
    (5) Any fuel-burning heater and other combustion equipment 
installation described in Sec. 29.859;
    (6) The compressor and accessory sections of turbine engines; and
    (7) The combustor, turbine, and tailpipe sections of turbine engine 
installations except sections that do not contain lines and components 
carrying flammable fluids or gases and are isolated from the designated 
fire zone prescribed in paragraph (a)(6) of this section by a firewall 
that meets Sec. 29.1191.
    (b) Each designated fire zone must meet the requirements of 
Sec. Sec. 29.1183 through 29.1203.

[Amdt. 29-3, 33 FR 970, Jan. 26, 1968, as amended by Amdt. 29-26, 53 FR 
34219, Sept. 2, 1988]



Sec. 29.1183  Lines, fittings, and components.

    (a) Except as provided in paragraph (b) of this section, each line, 
fitting, and other component carrying flammable fluid in any area 
subject to engine fire conditions and each component which conveys or 
contains flammable fluid in a designated fire zone must be fire 
resistant, except that flammable fluid tanks and supports in a 
designated fire zone must be fireproof or be enclosed by a fireproof 
shield unless damage by fire to any non-fireproof part will not cause 
leakage or spillage of flammable fluid. Components must be shielded or 
located so as to safeguard against the ignition of leaking flammable 
fluid. An integral oil sump of less than 25-quart capacity on a 
reciprocating engine need not be fireproof nor be enclosed by a 
fireproof shield.
    (b) Paragraph (a) of this section does not apply to--
    (1) Lines, fittings, and components which are already approved as 
part of a type certificated engine; and
    (2) Vent and drain lines, and their fittings, whose failure will not 
result in or add to, a fire hazard.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-2, 32 
FR 6914, May 5, 1967; Amdt. 29-10, 39 FR 35463, Oct. 1, 1974; Amdt. 29-
22, 49 FR 6850, Feb. 23, 1984]



Sec. 29.1185  Flammable fluids.

    (a) No tank or reservoir that is part of a system containing 
flammable fluids or gases may be in a designated fire zone unless the 
fluid contained, the design of the system, the materials used in the 
tank and its supports, the shutoff means, and the connections, lines, 
and controls provide a degree of safety equal to that which would exist 
if the tank or reservoir were outside such a zone.
    (b) Each fuel tank must be isolated from the engines by a firewall 
or shroud.
    (c) There must be at least one-half inch of clear airspace between 
each tank or reservoir and each firewall or shroud isolating a 
designated fire zone, unless equivalent means are used to prevent heat 
transfer from the fire zone to the flammable fluid.
    (d) Absorbent material close to flammable fluid system components 
that might leak must be covered or treated to prevent the absorption of 
hazardous quantities of fluids.



Sec. 29.1187  Drainage and ventilation of fire zones.

    (a) There must be complete drainage of each part of each designated 
fire zone to minimize the hazards resulting from failure or malfunction 
of any component containing flammable fluids. The drainage means must 
be--
    (1) Effective under conditions expected to prevail when drainage is 
needed; and

[[Page 821]]

    (2) Arranged so that no discharged fluid will cause an additional 
fire hazard.
    (b) Each designated fire zone must be ventilated to prevent the 
accumulation of flammable vapors.
    (c) No ventilation opening may be where it would allow the entry of 
flammable fluids, vapors, or flame from other zones.
    (d) Ventilation means must be arranged so that no discharged vapors 
will cause an additional fire hazard.
    (e) For category A rotorcraft, there must be means to allow the crew 
to shut off the sources of forced ventilation in any fire zone (other 
than the engine power section of the powerplant compartment) unless the 
amount of extinguishing agent and the rate of discharge are based on the 
maximum airflow through that zone.



Sec. 29.1189  Shutoff means.

    (a) There must be means to shut off or otherwise prevent hazardous 
quantities of fuel, oil, de-icing fluid, and other flammable fluids from 
flowing into, within, or through any designated fire zone, except that 
this means need not be provided--
    (1) For lines, fittings, and components forming an integral part of 
an engine;
    (2) For oil systems for turbine engine installations in which all 
components of the system, including oil tanks, are fireproof or located 
in areas not subject to engine fire conditions; or
    (3) For engine oil systems in category B rotorcraft using 
reciprocating engines of less than 500 cubic inches displacement.
    (b) The closing of any fuel shutoff valve for any engine may not 
make fuel unavailable to the remaining engines.
    (c) For category A rotorcraft, no hazardous quantity of flammable 
fluid may drain into any designated fire zone after shutoff has been 
accomplished, nor may the closing of any fuel shutoff valve for an 
engine make fuel unavailable to the remaining engines.
    (d) The operation of any shutoff may not interfere with the later 
emergency operation of any other equipment, such as the means for 
declutching the engine from the rotor drive.
    (e) Each shutoff valve and its control must be designed, located, 
and protected to function properly under any condition likely to result 
from fire in a designated fire zone.
    (f) Except for ground-use-only auxiliary power unit installations, 
there must be means to prevent inadvertent operation of each shutoff and 
to make it possible to reopen it in flight after it has been closed.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55473, Dec. 20, 1976; Amdt. 29-22, 49 FR 6850, Feb. 23, 1984; Amdt. 
29-26, 53 FR 34219, Sept. 2, 1988]



Sec. 29.1191  Firewalls.

    (a) Each engine, including the combustor, turbine, and tailpipe 
sections of turbine engine installations, must be isolated by a 
firewall, shroud, or equivalent means, from personnel compartments, 
structures, controls, rotor mechanisms, and other parts that are--
    (1) Essential to controlled flight and landing; and
    (2) Not protected under Sec. 29.861.
    (b) Each auxiliary power unit, combustion heater, and other 
combustion equipment to be used in flight, must be isolated from the 
rest of the rotorcraft by firewalls, shrouds, or equivalent means.
    (c) Each firewall or shroud must be constructed so that no hazardous 
quantity of air, fluid, or flame can pass from any engine compartment to 
other parts of the rotorcraft.
    (d) Each opening in the firewall or shroud must be sealed with 
close-fitting fireproof grommets, bushings, or firewall fittings.
    (e) Each firewall and shroud must be fireproof and protected against 
corrosion.
    (f) In meeting this section, account must be taken of the probable 
path of a fire as affected by the airflow in normal flight and in 
autorotation.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 970, Jan. 26, 1968]



Sec. 29.1193  Cowling and engine compartment covering.

    (a) Each cowling and engine compartment covering must be constructed 
and

[[Page 822]]

supported so that it can resist the vibration, inertia, and air loads to 
which it may be subjected in operation.
    (b) Cowling must meet the drainage and ventilation requirements of 
Sec. 29.1187.
    (c) On rotorcraft with a diaphragm isolating the engine power 
section from the engine accessory section, each part of the accessory 
section cowling subject to flame in case of fire in the engine power 
section of the powerplant must--
    (1) Be fireproof; and
    (2) Meet the requirements of Sec. 29.1191.
    (d) Each part of the cowling or engine compartment covering subject 
to high temperatures due to its nearness to exhaust system parts or 
exhaust gas impingement must be fireproof.
    (e) Each rotorcraft must--
    (1) Be designated and constructed so that no fire originating in any 
fire zone can enter, either through openings or by burning through 
external skin, any other zone or region where it would create additional 
hazards;
    (2) Meet the requirements of paragraph (e)(1) of this section with 
the landing gear retracted (if applicable); and
    (3) Have fireproof skin in areas subject to flame if a fire starts 
in or burns out of any designated fire zone.
    (f) A means of retention for each openable or readily removable 
panel, cowling, or engine or rotor drive system covering must be 
provided to preclude hazardous damage to rotors or critical control 
components in the event of--
    (1) Structural or mechanical failure of the normal retention means, 
unless such failure is extremely improbable; or
    (2) Fire in a fire zone, if such fire could adversely affect the 
normal means of retention.

(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 970, Jan. 26, 1968; Amdt. 29-13, 42 FR 15046, Mar. 17, 1977; Amdt. 
29-26, 53 FR 34219, Sept. 2, 1988]



Sec. 29.1194  Other surfaces.

    All surfaces aft of, and near, engine compartments and designated 
fire zones, other than tail surfaces not subject to heat, flames, or 
sparks emanating from a designated fire zone or engine compartment, must 
be at least fire resistant.

[Amdt. 29-3, 33 FR 970, Jan. 26, 1968]



Sec. 29.1195  Fire extinguishing systems.

    (a) Each turbine engine powered rotorcraft and Category A 
reciprocating engine powered rotorcraft, and each Category B 
reciprocating engine powered rotorcraft with engines of more than 1,500 
cubic inches must have a fire extinguishing system for the designated 
fire zones. The fire extinguishing system for a powerplant must be able 
to simultaneously protect all zones of the powerplant compartment for 
which protection is provided.
    (b) For multiengine powered rotorcraft, the fire extinguishing 
system, the quantity of extinguishing agent, and the rate of discharge 
must--
    (1) For each auxiliary power unit and combustion equipment, provide 
at least one adequate discharge; and
    (2) For each other designated fire zone, provide two adequate 
discharges.
    (c) For single engine rotorcraft, the quantity of extinguishing 
agent and the rate of discharge must provide at least one adequate 
discharge for the engine compartment.
    (d) It must be shown by either actual or simulated flight tests that 
under critical airflow conditions in flight the discharge of the 
extinguishing agent in each designated fire zone will provide an agent 
concentration capable of extinguishing fires in that zone and of 
minimizing the probability of reignition.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 
U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 970, Jan. 26, 1968; Amdt. 29-13, 42 FR 15047, Mar. 17, 1977; Amdt. 
29-17, 43 FR 50602, Oct. 30, 1978]



Sec. 29.1197  Fire extinguishing agents.

    (a) Fire extinguishing agents must--
    (1) Be capable of extinguishing flames emanating from any burning of 
fluids or other combustible materials

[[Page 823]]

in the area protected by the fire extinguishing system; and
    (2) Have thermal stability over the temperature range likely to be 
experienced in the compartment in which they are stored.
    (b) If any toxic extinguishing agent is used, it must be shown by 
test that entry of harmful concentrations of fluid or fluid vapors into 
any personnel compartment (due to leakage during normal operation of the 
rotorcraft, or discharge on the ground or in flight) is prevented, even 
though a defect may exist in the extinguishing system.

(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55473, Dec. 20, 1976; Amdt. 29-13, 42 FR 15047, Mar. 17, 1977]



Sec. 29.1199  Extinguishing agent containers.

    (a) Each extinguishing agent container must have a pressure relief 
to prevent bursting of the container by excessive internal pressures.
    (b) The discharge end of each discharge line from a pressure relief 
connection must be located so that discharge of the fire extinguishing 
agent would not damage the rotorcraft. The line must also be located or 
protected to prevent clogging caused by ice or other foreign matter.
    (c) There must be a means for each fire extinguishing agent 
container to indicate that the container has discharged or that the 
charging pressure is below the established minimum necessary for proper 
functioning.
    (d) The temperature of each container must be maintained, under 
intended operating conditions, to prevent the pressure in the container 
from--
    (1) Falling below that necessary to provide an adequate rate of 
discharge; or
    (2) Rising high enough to cause premature discharge.

(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655 (c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 
FR 15047, Mar. 17, 1977]



Sec. 29.1201  Fire extinguishing system materials.

    (a) No materials in any fire extinguishing system may react 
chemically with any extinguishing agent so as to create a hazard.
    (b) Each system component in an engine compartment must be 
fireproof.



Sec. 29.1203  Fire detector systems.

    (a) For each turbine engine powered rotorcraft and Category A 
reciprocating engine powered rotorcraft, and for each Category B 
reciprocating engine powered rotorcraft with engines of more than 900 
cubic inches displacement, there must be approved, quick-acting fire 
detectors in designated fire zones and in the combustor, turbine, and 
tailpipe sections of turbine installations (whether or not such sections 
are designated fire zones) in numbers and locations ensuring prompt 
detection of fire in those zones.
    (b) Each fire detector must be constructed and installed to 
withstand any vibration, inertia, and other loads to which it would be 
subjected in operation.
    (c) No fire detector may be affected by any oil, water, other 
fluids, or fumes that might be present.
    (d) There must be means to allow crewmembers to check, in flight, 
the functioning of each fire detector system electrical circuit.
    (e) The writing and other components of each fire detector system in 
an engine compartment must be at least fire resistant.
    (f) No fire detector system component for any fire zone may pass 
through another fire zone, unless--
    (1) It is protected against the possibility of false warnings 
resulting from fires in zones through which it passes; or
    (2) The zones involved are simultaneously protected by the same 
detector and extinguishing systems.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 970, Jan. 26, 1968]

[[Page 824]]



                           Subpart F_Equipment

                                 General



Sec. 29.1301  Function and installation.

    Each item of installed equipment must--
    (a) Be of a kind and design appropriate to its intended function;
    (b) Be labeled as to its identification, function, or operating 
limitations, or any applicable combination of these factors;
    (c) Be installed according to limitations specified for that 
equipment; and
    (d) Function properly when installed.



Sec. 29.1303  Flight and navigation instruments.

    The following are required flight and navigational instruments:
    (a) An airspeed indicator. For Category A rotorcraft with 
VNE less than a speed at which unmistakable pilot cues 
provide overspeed warning, a maximum allowable airspeed indicator must 
be provided. If maximum allowable airspeed varies with weight, altitude, 
temperature, or r.p.m., the indicator must show that variation.
    (b) A sensitive altimeter.
    (c) A magnetic direction indicator.
    (d) A clock displaying hours, minutes, and seconds with a sweep-
second pointer or digital presentation.
    (e) A free-air temperature indicator.
    (f) A non-tumbling gyroscopic bank and pitch indicator.
    (g) A gyroscopic rate-of-turn indicator combined with an integral 
slip-skid indicator (turn-and-bank indicator) except that only a slip-
skid indicator is required on rotorcraft with a third attitude 
instrument system that--
    (1) Is usable through flight attitudes of 80 
degrees of pitch and 120 degrees of roll;
    (2) Is powered from a source independent of the electrical 
generating system;
    (3) Continues reliable operation for a minimum of 30 minutes after 
total failure of the electrical generating system;
    (4) Operates independently of any other attitude indicating system;
    (5) Is operative without selection after total failure of the 
electrical generating system;
    (6) Is located on the instrument panel in a position acceptable to 
the Administrator that will make it plainly visible to and useable by 
any pilot at his station; and
    (7) Is appropriately lighted during all phases of operation.
    (h) A gyroscopic direction indicator.
    (i) A rate-of-climb (vertical speed) indicator.
    (j) For Category A rotorcraft, a speed warning device when 
VNE is less than the speed at which unmistakable overspeed 
warning is provided by other pilot cues. The speed warning device must 
give effective aural warning (differing distinctively from aural 
warnings used for other purposes) to the pilots whenever the indicated 
speed exceeds VNE plus 3 knots and must operate 
satisfactorily throughout the approved range of altitudes and 
temperatures.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55474, Dec. 20, 1976; Amdt. 29-14, 42 FR 36972, July 18, 1977; Amdt. 
29-24, 49 FR 44438, Nov. 6, 1984; 70 FR 2012, Jan. 12, 2005]



Sec. 29.1305  Powerplant instruments.

    The following are required powerplant instruments:
    (a) For each rotorcraft--
    (1) A carburetor air temperature indicator for each reciprocating 
engine;
    (2) A cylinder head temperature indicator for each air-cooled 
reciprocating engine, and a coolant temperature indicator for each 
liquid-cooled reciprocating engine;
    (3) A fuel quantity indicator for each fuel tank;
    (4) A low fuel warning device for each fuel tank which feeds an 
engine. This device must--
    (i) Provide a warning to the crew when approximately 10 minutes of 
usable fuel remains in the tank; and
    (ii) Be independent of the normal fuel quantity indicating system.
    (5) A manifold pressure indicator, for each reciprocating engine of 
the altitude type;
    (6) An oil pressure indicator for each pressure-lubricated gearbox.

[[Page 825]]

    (7) An oil pressure warning device for each pressure-lubricated 
gearbox to indicate when the oil pressure falls below a safe value;
    (8) An oil quantity indicator for each oil tank and each rotor drive 
gearbox, if lubricant is self-contained;
    (9) An oil temperature indicator for each engine;
    (10) An oil temperature warning device to indicate unsafe oil 
temperatures in each main rotor drive gearbox, including gearboxes 
necessary for rotor phasing;
    (11) A gas temperature indicator for each turbine engine;
    (12) A gas producer rotor tachometer for each turbine engine;
    (13) A tachometer for each engine that, if combined with the 
applicable instrument required by paragraph (a)(14) of this section, 
indicates rotor r.p.m. during autorotation.
    (14) At least one tachometer to indicate, as applicable--
    (i) The r.p.m. of the single main rotor;
    (ii) The common r.p.m. of any main rotors whose speeds cannot vary 
appreciably with respect to each other; and
    (iii) The r.p.m. of each main rotor whose speed can vary appreciably 
with respect to that of another main rotor;
    (15) A free power turbine tachometer for each turbine engine;
    (16) A means, for each turbine engine, to indicate power for that 
engine;
    (17) For each turbine engine, an indicator to indicate the 
functioning of the powerplant ice protection system;
    (18) An indicator for the filter required by Sec. 29.997 to 
indicate the occurrence of contamination of the filter to the degree 
established in compliance with Sec. 29.955;
    (19) For each turbine engine, a warning means for the oil strainer 
or filter required by Sec. 29.1019, if it has no bypass, to warn the 
pilot of the occurrence of contamination of the strainer or filter 
before it reaches the capacity established in accordance with Sec. 
29.1019(a)(2);
    (20) An indicator to indicate the functioning of any selectable or 
controllable heater used to prevent ice clogging of fuel system 
components;
    (21) An individual fuel pressure indicator for each engine, unless 
the fuel system which supplies that engine does not employ any pumps, 
filters, or other components subject to degradation or failure which may 
adversely affect fuel pressure at the engine;
    (22) A means to indicate to the flightcrew the failure of any fuel 
pump installed to show compliance with Sec. 29.955;
    (23) Warning or caution devices to signal to the flightcrew when 
ferromagnetic particles are detected by the chip detector required by 
Sec. 29.1337(e); and
    (24) For auxiliary power units, an individual indicator, warning or 
caution device, or other means to advise the flightcrew that limits are 
being exceeded, if exceeding these limits can be hazardous, for--
    (i) Gas temperature;
    (ii) Oil pressure; and
    (iii) Rotor speed.
    (25) For rotorcraft for which a 30-second/2-minute OEI power rating 
is requested, a means must be provided to alert the pilot when the 
engine is at the 30-second and 2-minute OEI power levels, when the event 
begins, and when the time interval expires.
    (26) For each turbine engine utilizing 30-second/2-minute OEI power, 
a device or system must be provided for use by ground personnel which--
    (i) Automatically records each usage and duration of power at the 
30-second and 2-minute OEI levels;
    (ii) Permits retrieval of the recorded data;
    (iii) Can be reset only by ground maintenance personnel; and
    (iv) Has a means to verify proper operation of the system or device.
    (b) For category A rotorcraft--
    (1) An individual oil pressure indicator for each engine, and either 
an independent warning device for each engine or a master warning device 
for the engines with means for isolating the individual warning circuit 
from the master warning device;
    (2) An independent fuel pressure warning device for each engine or a 
master warning device for all engines with provision for isolating the 
individual warning device from the master warning device; and
    (3) Fire warning indicators.
    (c) For category B rotorcraft--

[[Page 826]]

    (1) An individual oil pressure indicator for each engine; and
    (2) Fire warning indicators, when fire detection is required.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 970, Jan. 26, 1968; Amdt. 29-10, 39 FR 35463, Oct. 1, 1974; Amdt. 29-
26, 53 FR 34219, Sept. 2, 1988; Amdt. 29-34, 59 FR 47768, Sept. 16, 
1994; Amdt. 29-40, 61 FR 21908, May 10, 1996; 61 FR 43952, Aug. 27, 
1996]



Sec. 29.1307  Miscellaneous equipment.

    The following is required miscellaneous equipment:
    (a) An approved seat for each occupant.
    (b) A master switch arrangement for electrical circuits other than 
ignition.
    (c) Hand fire extinguishers.
    (d) A windshield wiper or equivalent device for each pilot station.
    (e) A two-way radio communication system.

[Amdt. 29-12, 41 FR 55473, Dec. 20, 1976]



Sec. 29.1309  Equipment, systems, and installations.

    (a) The equipment, systems, and installations whose functioning is 
required by this subchapter must be designed and installed to ensure 
that they perform their intended functions under any foreseeable 
operating condition.
    (b) The rotorcraft systems and associated components, considered 
separately and in relation to other systems, must be designed so that--
    (1) For Category B rotorcraft, the equipment, systems, and 
installations must be designed to prevent hazards to the rotorcraft if 
they malfunction or fail; or
    (2) For Category A rotorcraft--
    (i) The occurrence of any failure condition which would prevent the 
continued safe flight and landing of the rotorcraft is extremely 
improbable; and
    (ii) The occurrence of any other failure conditions which would 
reduce the capability of the rotorcraft or the ability of the crew to 
cope with adverse operating conditions is improbable.
    (c) Warning information must be provided to alert the crew to unsafe 
system operating conditions and to enable them to take appropriate 
corrective action. Systems, controls, and associated monitoring and 
warning means must be designed to minimize crew errors which could 
create additional hazards.
    (d) Compliance with the requirements of paragraph (b)(2) of this 
section must be shown by analysis and, where necessary, by appropriate 
ground, flight, or simulator tests. The analysis must consider--
    (1) Possible modes of failure, including malfunctions and damage 
from external sources;
    (2) The probability of multiple failures and undetected failures;
    (3) The resulting effects on the rotorcraft and occupants, 
considering the stage of flight and operating conditions; and
    (4) The crew warning cues, corrective action required, and the 
capability of detecting faults.
    (e) For Category A rotorcraft, each installation whose functioning 
is required by this subchapter and which requires a power supply is an 
``essential load'' on the power supply. The power sources and the system 
must be able to supply the following power loads in probable operating 
combinations and for probable durations:
    (1) Loads connected to the system with the system functioning 
normally.
    (2) Essential loads, after failure of any one prime mover, power 
converter, or energy storage device.
    (3) Essential loads, after failure of--
    (i) Any one engine, on rotorcraft with two engines; and
    (ii) Any two engines, on rotorcraft with three or more engines.
    (f) In determining compliance with paragraphs (e)(2) and (3) of this 
section, the power loads may be assumed to be reduced under a monitoring 
procedure consistent with safety in the kinds of operations authorized. 
Loads not required for controlled flight need not be considered for the 
two-engine-inoperative condition on rotorcraft with three or more 
engines.
    (g) In showing compliance with paragraphs (a) and (b) of this 
section with regard to the electrical system and to equipment design and 
installation, critical environmental conditions must be considered. For 
electrical generation, distribution, and utilization

[[Page 827]]

equipment required by or used in complying with this subchapter, except 
equipment covered by Technical Standard Orders containing environmental 
test procedures, the ability to provide continuous, safe service under 
foreseeable environmental conditions may be shown by environmental 
tests, design analysis, or reference to previous comparable service 
experience on other aircraft.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 
FR 36972, July 18, 1977; Amdt. 29-24, 49 FR 44438, Nov. 6, 1984; Amdt. 
29-40, 61 FR 21908, May 10, 1996; Amdt. 29-53, 76 FR 33136, June 8, 
2011]



Sec. 29.1316  Electrical and electronic system lightning protection.

    (a) Each electrical and electronic system that performs a function, 
for which failure would prevent the continued safe flight and landing of 
the rotorcraft, must be designed and installed so that--
    (1) The function is not adversely affected during and after the time 
the rotorcraft is exposed to lightning; and
    (2) The system automatically recovers normal operation of that 
function in a timely manner after the rotorcraft is exposed to 
lightning.
    (b) Each electrical and electronic system that performs a function, 
for which failure would reduce the capability of the rotorcraft or the 
ability of the flightcrew to respond to an adverse operating condition, 
must be designed and installed so that the function recovers normal 
operation in a timely manner after the rotorcraft is exposed to 
lightning.

[Doc. No. FAA-2010-0224, Amdt. 29-53, 76 FR 33136, June 8, 2011]



Sec. 29.1317  High-intensity Radiated Fields (HIRF) Protection.

    (a) Except as provided in paragraph (d) of this section, each 
electrical and electronic system that performs a function whose failure 
would prevent the continued safe flight and landing of the rotorcraft 
must be designed and installed so that--
    (1) The function is not adversely affected during and after the time 
the rotorcraft is exposed to HIRF environment I, as described in 
appendix E to this part;
    (2) The system automatically recovers normal operation of that 
function, in a timely manner, after the rotorcraft is exposed to HIRF 
environment I, as described in appendix E to this part, unless this 
conflicts with other operational or functional requirements of that 
system;
    (3) The system is not adversely affected during and after the time 
the rotorcraft is exposed to HIRF environment II, as described in 
appendix E to this part; and
    (4) Each function required during operation under visual flight 
rules is not adversely affected during and after the time the rotorcraft 
is exposed to HIRF environment III, as described in appendix E to this 
part.
    (b) Each electrical and electronic system that performs a function 
whose failure would significantly reduce the capability of the 
rotorcraft or the ability of the flightcrew to respond to an adverse 
operating condition must be designed and installed so the system is not 
adversely affected when the equipment providing these functions is 
exposed to equipment HIRF test level 1 or 2, as described in appendix E 
to this part.
    (c) Each electrical and electronic system that performs such a 
function whose failure would reduce the capability of the rotorcraft or 
the ability of the flightcrew to respond to an adverse operating 
condition must be designed and installed so the system is not adversely 
affected when the equipment providing these functions is exposed to 
equipment HIRF test level 3, as described in appendix E to this part.
    (d) Before December 1, 2012, an electrical or electronic system that 
performs a function whose failure would prevent the continued safe 
flight and landing of a rotorcraft may be designed and installed without 
meeting the provisions of paragraph (a) provided--
    (1) The system has previously been shown to comply with special 
conditions for HIRF, prescribed under Sec. 21.16, issued before 
December 1, 2007;

[[Page 828]]

    (2) The HIRF immunity characteristics of the system have not changed 
since compliance with the special conditions was demonstrated; and
    (3) The data used to demonstrate compliance with the special 
conditions is provided.

[Doc. No. FAA-2006-23657, 72 FR 44027, Aug. 6, 2007]

                        Instruments: Installation



Sec. 29.1321  Arrangement and visibility.

    (a) Each flight, navigation, and powerplant instrument for use by 
any pilot must be easily visible to him from his station with the 
minimum practicable deviation from his normal position and line of 
vision when he is looking forward along the flight path.
    (b) Each instrument necessary for safe operation, including the 
airspeed indicator, gyroscopic direction indicator, gyroscopic bank-and-
pitch indicator, slip-skid indicator, altimeter, rate-of-climb 
indicator, rotor tachometers, and the indicator most representative of 
engine power, must be grouped and centered as nearly as practicable 
about the vertical plane of the pilot's forward vision. In addition, for 
rotorcraft approved for IFR flight--
    (1) The instrument that most effectively indicates attitude must be 
on the panel in the top center position;
    (2) The instrument that most effectively indicates direction of 
flight must be adjacent to and directly below the attitude instrument;
    (3) The instrument that most effectively indicates airspeed must be 
adjacent to and to the left of the attitude instrument; and
    (4) The instrument that most effectively indicates altitude or is 
most frequently utilized in control of altitude must be adjacent to and 
to the right of the attitude instrument.
    (c) Other required powerplant instruments must be closely grouped on 
the instrument panel.
    (d) Identical powerplant instruments for the engines must be located 
so as to prevent any confusion as to which engine each instrument 
relates.
    (e) Each powerplant instrument vital to safe operation must be 
plainly visible to appropriate crewmembers.
    (f) Instrument panel vibration may not damage, or impair the 
readability or accuracy of, any instrument.
    (g) If a visual indicator is provided to indicate malfunction of an 
instrument, it must be effective under all probable cockpit lighting 
conditions.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 
FR 36972, July 18, 1977; Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]



Sec. 29.1322  Warning, caution, and advisory lights.

    If warning, caution or advisory lights are installed in the cockpit 
they must, unless otherwise approved by the Administrator, be--
    (a) Red, for warning lights (lights indicating a hazard which may 
require immediate corrective action);
    (b) Amber, for caution lights (lights indicating the possible need 
for future corrective action);
    (c) Green, for safe operation lights; and
    (d) Any other color, including white, for lights not described in 
paragraphs (a) through (c) of this section, provided the color differs 
sufficiently from the colors prescribed in paragraphs (a) through (c) of 
this section to avoid possible confusion.

[Amdt. 29-12, 41 FR 55474, Dec. 20, 1976]



Sec. 29.1323  Airspeed indicating system.

    For each airspeed indicating system, the following apply:
    (a) Each airspeed indicating instrument must be calibrated to 
indicate true airspeed (at sea level with a standard atmosphere) with a 
minimum practicable instrument calibration error when the corresponding 
pitot and static pressures are applied.
    (b) Each system must be calibrated to determine system error 
excluding airspeed instrument error. This calibration must be 
determined--
    (1) In level flight at speeds of 20 knots and greater, and over an 
appropriate range of speeds for flight conditions of climb and 
autorotation; and
    (2) During takeoff, with repeatable and readable indications that 
ensure--

[[Page 829]]

    (i) Consistent realization of the field lengths specified in the 
Rotorcraft Flight Manual; and
    (ii) Avoidance of the critical areas of the height-velocity envelope 
as established under Sec. 29.87.
    (c) For Category A rotorcraft--
    (1) The indication must allow consistent definition of the takeoff 
decision point; and
    (2) The system error, excluding the airspeed instrument calibration 
error, may not exceed--
    (i) Three percent or 5 knots, whichever is greater, in level flight 
at speeds above 80 percent of takeoff safety speed; and
    (ii) Ten knots in climb at speeds from 10 knots below takeoff safety 
speed to 10 knots above VY.
    (d) For Category B rotorcraft, the system error, excluding the 
airspeed instrument calibration error, may not exceed 3 percent or 5 
knots, whichever is greater, in level flight at speeds above 80 percent 
of the climbout speed attained at 50 feet when complying with Sec. 
29.63.
    (e) Each system must be arranged, so far as practicable, to prevent 
malfunction or serious error due to the entry of moisture, dirt, or 
other substances.
    (f) Each system must have a heated pitot tube or an equivalent means 
of preventing malfunction due to icing.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964 as amended by Amdt. 29-3, 33 
FR 970, Jan. 26, 1968; Amdt. 29-24, 49 FR 44439, Nov. 6, 1984; Amdt. 29-
39, 61 FR 21901, May 10, 1996; Amdt. 29-44, 64 FR 45338, Aug. 19, 1999]



Sec. 29.1325  Static pressure and pressure altimeter systems.

    (a) Each instrument with static air case connections must be vented 
to the outside atmosphere through an appropriate piping system.
    (b) Each vent must be located where its orifices are least affected 
by airflow variation, moisture, or foreign matter.
    (c) Each static pressure port must be designed and located in such 
manner that the correlation between air pressure in the static pressure 
system and true ambient atmospheric static pressure is not altered when 
the rotorcraft encounters icing conditions. An anti-icing means or an 
alternate source of static pressure may be used in showing compliance 
with this requirement. If the reading of the altimeter, when on the 
alternate static pressure system, differs from the reading of altimeter 
when on the primary static system by more than 50 feet, a correction 
card must be provided for the alternate static system.
    (d) Except for the vent into the atmosphere, each system must be 
airtight.
    (e) Each pressure altimeter must be approved and calibrated to 
indicate pressure altitude in a standard atmosphere with a minimum 
practicable calibration error when the corresponding static pressures 
are applied.
    (f) Each system must be designed and installed so that an error in 
indicated pressure altitude, at sea level, with a standard atmosphere, 
excluding instrument calibration error, does not result in an error of 
more than 30 feet per 100 knots speed. However, 
the error need not be less than 30 feet.
    (g) Except as provided in paragraph (h) of this section, if the 
static pressure system incorporates both a primary and an alternate 
static pressure source, the means for selecting one or the other source 
must be designed so that--
    (1) When either source is selected, the other is blocked off; and
    (2) Both sources cannot be blocked off simultaneously.
    (h) For unpressurized rotorcraft, paragraph (g)(1) of this section 
does not apply if it can be demonstrated that the static pressure system 
calibration, when either static pressure source is selected, is not 
changed by the other static pressure source being open or blocked.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 
FR 36972, July 18, 1977; Amdt. 29-24, 49 FR 44439, Nov. 6, 1984]



Sec. 29.1327  Magnetic direction indicator.

    (a) Each magnetic direction indicator must be installed so that its 
accuracy is not excessively affected by the rotor
craft's vibration or magnetic fields.

[[Page 830]]

    (b) The compensated installation may not have a deviation, in level 
flight, greater than 10 degrees on any heading.



Sec. 29.1329  Automatic pilot system.

    (a) Each automatic pilot system must be designed so that the 
automatic pilot can--
    (1) Be sufficiently overpowered by one pilot to allow control of the 
rotorcraft; and
    (2) Be readily and positively disengaged by each pilot to prevent it 
from interfering with the control of the rotorcraft.
    (b) Unless there is automatic synchronization, each system must have 
a means to readily indicate to the pilot the alignment of the actuating 
device in relation to the control system it operates.
    (c) Each manually operated control for the system's operation must 
be readily accessible to the pilots.
    (d) The system must be designed and adjusted so that, within the 
range of adjustment available to the pilot, it cannot produce hazardous 
loads on the rotorcraft, or create hazardous deviations in the flight 
path, under any flight condition appropriate to its use, either during 
normal operation or in the event of a malfunction, assuming that 
corrective action begins within a reasonable period of time.
    (e) If the automatic pilot integrates signals from auxiliary 
controls or furnishes signals for operation of other equipment, there 
must be positive interlocks and sequencing of engagement to prevent 
improper operation.
    (f) If the automatic pilot system can be coupled to airborne 
navigation equipment, means must be provided to indicate to the pilots 
the current mode of operation. Selector switch position is not 
acceptable as a means of indication.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44439, Nov. 6, 1984; Amdt. 29-24, 49 FR 47594, Dec. 6, 1984; Amdt. 
29-42, 63 FR 43285, Aug. 12, 1998]



Sec. 29.1331  Instruments using a power supply.

    For category A rotorcraft--
    (a) Each required flight instrument using a power supply must have--
    (1) Two independent sources of power;
    (2) A means of selecting either power source; and
    (3) A visual means integral with each instrument to indicate when 
the power adequate to sustain proper instrument performance is not being 
supplied. The power must be measured at or near the point where it 
enters the instrument. For electrical instruments, the power is 
considered to be adequate when the voltage is within the approved 
limits; and
    (b) The installation and power supply system must be such that 
failure of any flight instrument connected to one source, or of the 
energy supply from one source, or a fault in any part of the power 
distribution system does not interfere with the proper supply of energy 
from any other source.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44439, Nov. 6, 1984]



Sec. 29.1333  Instrument systems.

    For systems that operate the required flight instruments which are 
located at each pilot's station, the following apply:
    (a) Only the required flight instruments for the first pilot may be 
connected to that operating system.
    (b) The equipment, systems, and installations must be designed so 
that one display of the information essential to the safety of flight 
which is provided by the flight instruments remains available to a 
pilot, without additional crewmember action, after any single failure or 
combination of failures that are not shown to be extremely improbable.
    (c) Additional instruments, systems, or equipment may not be 
connected to the operating system for a second pilot unless provisions 
are made to ensure the continued normal functioning of the required 
flight instruments in the event of any malfunction of the additional 
instruments, systems, or equipment which is not shown to be extremely 
improbable.

[Amdt. 29-24, 49 FR 44439, Nov. 6, 1984]

[[Page 831]]



Sec. 29.1335  Flight director systems.

    If a flight director system is installed, means must be provided to 
indicate to the flight crew its current mode of operation. Selector 
switch position is not acceptable as a means of indication.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 29-14, 42 FR 36973, July 18, 1977]



Sec. 29.1337  Powerplant instruments.

    (a) Instruments and instrument lines. (1) Each powerplant and 
auxiliary power unit instrument line must meet the requirements of 
Sec. Sec. 29.993 and 29.1183.
    (2) Each line carrying flammable fluids under pressure must--
    (i) Have restricting orifices or other safety devices at the source 
of pressure to prevent the escape of excessive fluid if the line fails; 
and
    (ii) Be installed and located so that the escape of fluids would not 
create a hazard.
    (3) Each powerplant and auxiliary power unit instrument that 
utilizes flammable fluids must be installed and located so that the 
escape of fluid would not create a hazard.
    (b) Fuel quantity indicator. There must be means to indicate to the 
flight crew members the quantity, in gallons or equivalent units, of 
usable fuel in each tank during flight. In addition--
    (1) Each fuel quantity indicator must be calibrated to read ``zero'' 
during level flight when the quantity of fuel remaining in the tank is 
equal to the unusable fuel supply determined under Sec. 29.959;
    (2) When two or more tanks are closely interconnected by a gravity 
feed system and vented, and when it is impossible to feed from each tank 
separately, at least one fuel quantity indicator must be installed;
    (3) Tanks with interconnected outlets and airspaces may be treated 
as one tank and need not have separate indicators; and
    (4) Each exposed sight gauge used as a fuel quantity indicator must 
be protected against damage.
    (c) Fuel flowmeter system. If a fuel flowmeter system is installed, 
each metering component must have a means for bypassing the fuel supply 
if malfunction of that component severely restricts fuel flow.
    (d) Oil quantity indicator. There must be a stick gauge or 
equivalent means to indicate the quantity of oil--
    (1) In each tank; and
    (2) In each transmission gearbox.
    (e) Rotor drive system transmissions and gearboxes utilizing 
ferromagnetic materials must be equipped with chip detectors designed to 
indicate the presence of ferromagnetic particles resulting from damage 
or excessive wear within the transmission or gearbox. Each chip detector 
must--
    (1) Be designed to provide a signal to the indicator required by 
Sec. 29.1305(a)(22); and
    (2) Be provided with a means to allow crewmembers to check, in 
flight, the function of each detector electrical circuit and signal.

(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 
FR 15047, Mar. 17, 1977; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]

                    Electrical Systems and Equipment



Sec. 29.1351  General.

    (a) Electrical system capacity. The required generating capacity and 
the number and kind of power sources must--
    (1) Be determined by an electrical load analysis; and
    (2) Meet the requirements of Sec. 29.1309.
    (b) Generating system. The generating system includes electrical 
power sources, main power busses, transmission cables, and associated 
control, regulation, and protective devices. It must be designed so 
that--
    (1) Power sources function properly when independent and when 
connected in combination;
    (2) No failure or malfunction of any power source can create a 
hazard or impair the ability of remaining sources to supply essential 
loads;
    (3) The system voltage and frequency (as applicable) at the 
terminals of essential load equipment can be maintained within the 
limits for which the

[[Page 832]]

equipment is designed, during any probable operating condition;
    (4) System transients due to switching, fault clearing, or other 
causes do not make essential loads inoperative, and do not cause a smoke 
or fire hazard;
    (5) There are means accessible in flight to appropriate crewmembers 
for the individual and collective disconnection of the electrical power 
sources from the main bus; and
    (6) There are means to indicate to appropriate crewmembers the 
generating system quantities essential for the safe operation of the 
system, such as the voltage and current supplied by each generator.
    (c) External power. If provisions are made for connecting external 
power to the rotorcraft, and that external power can be electrically 
connected to equipment other than that used for engine starting, means 
must be provided to ensure that no external power supply having a 
reverse polarity, or a reverse phase sequence, can supply power to the 
rotorcraft's electrical system.
    (d) Operation with the normal electrical power generating system 
inoperative.
    (1) It must be shown by analysis, tests, or both, that the 
rotorcraft can be operated safely in VFR conditions for a period of not 
less than 5 minutes, with the normal electrical power generating system 
(electrical power sources excluding the battery) inoperative, with 
critical type fuel (from the standpoint of flameout and restart 
capability), and with the rotorcraft initially at the maximum 
certificated altitude. Parts of the electrical system may remain on if--
    (i) A single malfunction, including a wire bundle or junction box 
fire, cannot result in loss of the part turned off and the part turned 
on;
    (ii) The parts turned on are electrically and mechanically isolated 
from the parts turned off; and
    (2) Additional requirements for Category A Rotorcraft.
    (i) Unless it can be shown that the loss of the normal electrical 
power generating system is extremely improbable, an emergency electrical 
power system, independent of the normal electrical power generating 
system, must be provided, with sufficient capacity to power all systems 
necessary for continued safe flight and landing.
    (ii) Failures, including junction box, control panel, or wire bundle 
fires, which would result in the loss of the normal and emergency 
systems, must be shown to be extremely improbable.
    (iii) Systems necessary for immediate safety must continue to 
operate following the loss of the normal electrical power generating 
system, without the need for flight crew action.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 
FR 36973, July 18, 1977; Amdt. 29-40, 61 FR 21908, May 10, 1996; Amdt. 
29-42, 63 FR 43285, Aug. 12, 1998]



Sec. 29.1353  Electrical equipment and installations.

    (a) Electrical equipment, controls, and wiring must be installed so 
that operation of any one unit or system of units will not adversely 
affect the simultaneous operation of any other electrical unit or system 
essential to safe operation.
    (b) Cables must be grouped, routed, and spaced so that damage to 
essential circuits will be minimized if there are faults in heavy 
current-carrying cables.
    (c) Storage batteries must be designed and installed as follows:
    (1) Safe cell temperatures and pressures must be maintained during 
any probable charging and discharging condition. No uncontrolled 
increase in cell temperature may result when the battery is recharged 
(after previous complete discharge)--
    (i) At maximum regulated voltage or power;
    (ii) During a flight of maximum duration; and
    (iii) Under the most adverse cooling condition likely in service.
    (2) Compliance with paragraph (a)(1) of this section must be shown 
by test unless experience with similar batteries and installations has 
shown that maintaining safe cell temperatures and pressures presents no 
problem.

[[Page 833]]

    (3) No explosive or toxic gases emitted by any battery in normal 
operation, or as the result of any probable malfunction in the charging 
system or battery installation, may accumulate in hazardous quantities 
within the rotorcraft.
    (4) No corrosive fluids or gases that may escape from the battery 
may damage surrounding structures or adjacent essential equipment.
    (5) Each nickel cadmium battery installation capable of being used 
to start an engine or auxiliary power unit must have provisions to 
prevent any hazardous effect on structure or essential systems that may 
be caused by the maximum amount of heat the battery can generate during 
a short circuit of the battery or of its individual cells.
    (6) Nickel cadmium battery installations capable of being used to 
start an engine or auxiliary power unit must have--
    (i) A system to control the charging rate of the battery 
automatically so as to prevent battery overheating;
    (ii) A battery temperature sensing and over-temperature warning 
system with a means for disconnecting the battery from its charging 
source in the event of an over-temperature condition; or
    (iii) A battery failure sensing and warning system with a means for 
disconnecting the battery from its charging source in the event of 
battery failure.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 
FR 36973, July 18, 1977; Amdt. 29-15, 43 FR 2327, Jan. 16, 1978]



Sec. 29.1355  Distribution system.

    (a) The distribution system includes the distribution busses, their 
associated feeders, and each control and protective device.
    (b) If two independent sources of electrical power for particular 
equipment or systems are required by this chapter, in the event of the 
failure of one power source for such equipment or system, another power 
source (including its separate feeder) must be provided automatically or 
be manually selectable to maintain equipment or system operation.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 
FR 36973, July 18, 1977; Amdt. 29-24, 49 FR 44439, Nov. 6, 1984]



Sec. 29.1357  Circuit protective devices.

    (a) Automatic protective devices must be used to minimize distress 
to the electrical system and hazard to the rotorcraft system and hazard 
to the rotorcraft in the event of wiring faults or serious malfunction 
of the system or connected equipment.
    (b) The protective and control devices in the generating system must 
be designed to de-energize and disconnect faulty power sources and power 
transmission equipment from their associated buses with sufficient 
rapidity to provide protection from hazardous overvoltage and other 
malfunctioning.
    (c) Each resettable circuit protective device must be designed so 
that, when an overload or circuit fault exists, it will open the circuit 
regardless of the position of the operating control.
    (d) If the ability to reset a circuit breaker or replace a fuse is 
essential to safety in flight, that circuit breaker or fuse must be 
located and identified so that it can be readily reset or replaced in 
flight.
    (e) Each essential load must have individual circuit protection. 
However, individual protection for each circuit in an essential load 
system (such as each position light circuit in a system) is not 
required.
    (f) If fuses are used, there must be spare fuses for use in flight 
equal to at least 50 percent of the number of fuses of each rating 
required for complete circuit protection.
    (g) Automatic reset circuit breakers may be used as integral 
protectors for electrical equipment provided there is circuit protection 
for the cable supplying power to the equipment.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 
FR 44440, Nov. 6, 1984]

[[Page 834]]



Sec. 29.1359  Electrical system fire and smoke protection.

    (a) Components of the electrical system must meet the applicable 
fire and smoke protection provisions of Sec. Sec. 29.831 and 29.863.
    (b) Electrical cables, terminals, and equipment, in designated fire 
zones, and that are used in emergency procedures, must be at least fire 
resistant.
    (c) Insulation on electrical wire and cable installed in the 
rotorcraft must be self-extinguishing when tested in accordance with 
Appendix F, Part I(a)(3), of part 25 of this chapter.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-42, 63 
FR 43285, Aug. 12, 1998]



Sec. 29.1363  Electrical system tests.

    (a) When laboratory tests of the electrical system are conducted--
    (1) The tests must be performed on a mock-up using the same 
generating equipment used in the rotorcraft;
    (2) The equipment must simulate the electrical characteristics of 
the distribution wiring and connected loads to the extent necessary for 
valid test results; and
    (3) Laboratory generator drives must simulate the prime movers on 
the rotorcraft with respect to their reaction to generator loading, 
including loading due to faults.
    (b) For each flight condition that cannot be simulated adequately in 
the laboratory or by ground tests on the rotorcraft, flight tests must 
be made.

                                 Lights



Sec. 29.1381  Instrument lights.

    The instrument lights must--
    (a) Make each instrument, switch, and other device for which they 
are provided easily readable; and
    (b) Be installed so that--
    (1) Their direct rays are shielded from the pilot's eyes; and
    (2) No objectionable reflections are visible to the pilot.



Sec. 29.1383  Landing lights.

    (a) Each required landing or hovering light must be approved.
    (b) Each landing light must be installed so that--
    (1) No objectionable glare is visible to the pilot;
    (2) The pilot is not adversely affected by halation; and
    (3) It provides enough light for night operation, including hovering 
and landing.
    (c) At least one separate switch must be provided, as applicable--
    (1) For each separately installed landing light; and
    (2) For each group of landing lights installed at a common location.



Sec. 29.1385  Position light system installation.

    (a) General. Each part of each position light system must meet the 
applicable requirements of this section and each system as a whole must 
meet the requirements of Sec. Sec. 29.1387 through 29.1397.
    (b) Forward position lights. Forward position lights must consist of 
a red and a green light spaced laterally as far apart as practicable and 
installed forward on the rotorcraft so that, with the rotorcraft in the 
normal flying position, the red light is on the left side, and the green 
light is on the right side. Each light must be approved.
    (c) Rear position light. The rear position light must be a white 
light mounted as far aft as practicable, and must be approved.
    (d) Circuit. The two forward position lights and the rear position 
light must make a single circuit.
    (e) Light covers and color filters. Each light cover or color filter 
must be at least flame resistant and may not change color or shape or 
lose any appreciable light transmission during normal use.



Sec. 29.1387  Position light system dihedral angles.

    (a) Except as provided in paragraph (e) of this section, each 
forward and rear position light must, as installed, show unbroken light 
within the dihedral angles described in this section.
    (b) Dihedral angle L (left) is formed by two intersecting vertical 
planes, the first parallel to the longitudinal axis of the rotorcraft, 
and the other at 110 degrees to the left of the first, as viewed when 
looking forward along the longitudinal axis.

[[Page 835]]

    (c) Dihedral angle R (right) is formed by two intersecting vertical 
planes, the first parallel to the longitudinal axis of the rotorcraft, 
and the other at 110 degrees to the right of the first, as viewed when 
looking forward along the longitudinal axis.
    (d) Dihedral angle A (aft) is formed by two intersecting vertical 
planes making angles of 70 degrees to the right and to the left, 
respectively, to a vertical plane passing through the longitudinal axis, 
as viewed when looking aft along the longitudinal axis.
    (e) If the rear position light, when mounted as far aft as 
practicable in accordance with Sec. 29.1385(c), cannot show unbroken 
light within dihedral angle A (as defined in paragraph (d) of this 
section), a solid angle or angles of obstructed visibility totaling not 
more than 0.04 steradians is allowable within that dihedral angle, if 
such solid angle is within a cone whose apex is at the rear position 
light and whose elements make an angle of 30[deg] with a vertical line 
passing through the rear position light.

(49 U.S.C. 1655(c))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-9, 36 
FR 21279, Nov. 5, 1971]



Sec. 29.1389  Position light distribution and intensities.

    (a) General. The intensities prescribed in this section must be 
provided by new equipment with light covers and color filters in place. 
Intensities must be determined with the light source operating at a 
steady value equal to the average luminous output of the source at the 
normal operating voltage of the rotorcraft. The light distribution and 
intensity of each position light must meet the requirements of paragraph 
(b) of this section.
    (b) Forward and rear position lights. The light distribution and 
intensities of forward and rear position lights must be expressed in 
terms of minimum intensities in the horizontal plane, minimum 
intensities in any vertical plane, and maximum intensities in 
overlapping beams, within dihedral angles, L, R, and A, and must meet 
the following requirements:
    (1) Intensities in the horizontal plane. Each intensity in the 
horizontal plane (the plane containing the longitudinal axis of the 
rotorcraft and perpendicular to the plane of symmetry of the 
rotorcraft), must equal or exceed the values in Sec. 29.1391.
    (2) Intensities in any vertical plane. Each intensity in any 
vertical plane (the plane perpendicular to the horizontal plane) must 
equal or exceed the appropriate value in Sec. 29.1393 where I is the 
minimum intensity prescribed in Sec. 29.1391 for the corresponding 
angles in the horizontal plane.
    (3) Intensities in overlaps between adjacent signals. No intensity 
in any overlap between adjacent signals may exceed the values in Sec. 
29.1395, except that higher intensities in overlaps may be used with the 
use of main beam intensities substantially greater than the minima 
specified in Sec. Sec. 29.1391 and 29.1393 if the overlap intensities 
in relation to the main beam intensities do not adversely affect signal 
clarity.



Sec. 29.1391  Minimum intensities in the horizontal plane of forward
and rear position lights.

    Each position light intensity must equal or exceed the applicable 
values in the following table:

------------------------------------------------------------------------
                                        Angle from right or
                                       left of longitudinal    Intensity
   Dihedral angle (light included)      axis, measured from    (candles)
                                            dead ahead
------------------------------------------------------------------------
L and R (forward red and green).....  0[deg] to 10[deg].....          40
                                      10[deg] to 20[deg]....          30
                                      20[deg] to 110[deg]...           5
A (rear white)......................  110[deg] to 180[deg]..          20
------------------------------------------------------------------------



Sec. 29.1393  Minimum intensities in any vertical plane of forward
and rear position lights.

    Each position light intensity must equal or exceed the applicable 
values in the following table:

------------------------------------------------------------------------
                                                              Intensity,
          Angle above or below the horizontal plane                I
------------------------------------------------------------------------
0[deg]......................................................        1.00
0[deg] to 5[deg]............................................         .90
5[deg] to 10[deg]...........................................         .80
10[deg] to 15[deg]..........................................         .70
15[deg] to 20[deg]..........................................         .50
20[deg] to 30[deg]..........................................         .30
30[deg] to 40[deg]..........................................         .10
40[deg] to 90[deg]..........................................         .05
------------------------------------------------------------------------


[[Page 836]]



Sec. 29.1395  Maximum intensities in overlapping beams of forward 
and rear position lights.

    No position light intensity may exceed the applicable values in the 
following table, except as provided in Sec. 29.1389(b)(3).

------------------------------------------------------------------------
                                                    Maximum intensity
                                               -------------------------
                   Overlaps                        Area A       Area B
                                                 (candles)    (candles)
------------------------------------------------------------------------
Green in dihedral angle L.....................           10            1
Red in dihedral angle R.......................           10            1
Green in dihedral angle A.....................            5            1
Red in dihedral angle A.......................            5            1
Rear white in dihedral angle L................            5            1
Rear white in dihedral angle R................            5            1
------------------------------------------------------------------------


Where--
    (a) Area A includes all directions in the adjacent dihedral angle 
that pass through the light source and intersect the common boundary 
plane at more than 10 degrees but less than 20 degrees; and
    (b) Area B includes all directions in the adjacent dihedral angle 
that pass through the light source and intersect the common boundary 
plane at more than 20 degrees.



Sec. 29.1397  Color specifications.

    Each position light color must have the applicable International 
Commission on Illumination chromaticity coordinates as follows:
    (a) Aviation red--

    y is not greater than 0.335; and
    z is not greater than 0.002.

    (b) Aviation green--

    x is not greater than 0.440-0.320y;
    x is not greater than y-0.170; and
    y is not less than 0.390-0.170x.

    (c) Aviation white--

    x is not less than 0.300 and not greater than 0.540;
    y is not less than x-0.040 or yc-0.010, whichever is the smaller; 
and
    y is not greater than x+0.020 nor 0.636-0.400x;
    Where Ye is the y coordinate of the Planckian radiator for the value 
of x considered.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-7, 36 
FR 12972, July 10, 1971]



Sec. 29.1399  Riding light.

    (a) Each riding light required for water operation must be installed 
so that it can--
    (1) Show a white light for at least two miles at night under clear 
atmospheric conditions; and
    (2) Show a maximum practicable unbroken light with the rotorcraft on 
the water.
    (b) Externally hung lights may be used.



Sec. 29.1401  Anticollision light system.

    (a) General. If certification for night operation is requested, the 
rotorcraft must have an anticollision light system that--
    (1) Consists of one or more approved anticollision lights located so 
that their emitted light will not impair the crew's vision or detract 
from the conspicuity of the position lights; and
    (2) Meets the requirements of paragraphs (b) through (f) of this 
section.
    (b) Field of coverage. The system must consist of enough lights to 
illuminate the vital areas around the rotorcraft, considering the 
physical configuration and flight characteristics of the rotorcraft. The 
field of coverage must extend in each direction within at least 30 
degrees above and 30 degrees below the horizontal plane of the 
rotorcraft, except that there may be solid angles of obstructed 
visibility totaling not more than 0.5 steradians.
    (c) Flashing characteristics. The arrangement of the system, that 
is, the number of light sources, beam width, speed of rotation, and 
other characteristics, must give an effective flash frequency of not 
less than 40, nor more than 100, cycles per minute. The effective flash 
frequency is the frequency at which the rotorcraft's complete 
anticollision light system is observed from a distance, and applies to 
each sector of light including any overlaps that exist when the system 
consists of more than one light source. In overlaps, flash frequencies 
may exceed 100, but not 180, cycles per minute.
    (d) Color. Each anticollision light must be aviation red and must 
meet the applicable requirements of Sec. 29.1397.
    (e) Light intensity. The minimum light intensities in any vertical 
plane, measured with the red filter (if used)

[[Page 837]]

and expressed in terms of ``effective'' intensities must meet the 
requirements of paragraph (f) of this section. The following relation 
must be assumed:
[GRAPHIC] [TIFF OMITTED] TC28SE91.090

where:

Ie=effective intensity (candles).
I(t)=instantaneous intensity as a function of time.
t2-tl=flash time interval (seconds).
Normally, the maximum value of effective intensity is obtained when t2 
and t1 are chosen so that the effective intensity is equal to the 
instantaneous intensity at t2 and t1.

    (f) Minimum effective intensities for anticollision light. Each 
anticollision light effective intensity must equal or exceed the 
applicable values in the following table:

------------------------------------------------------------------------
                                                               Effective
          Angle above or below the horizontal plane            intensity
                                                               (candles)
------------------------------------------------------------------------
0[deg] to 5[deg]............................................         150
5[deg] to 10[deg]...........................................          90
10[deg] to 20[deg]..........................................          30
20[deg] to 30[deg]..........................................          15
------------------------------------------------------------------------


[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-7, 36 
FR 12972, July 10, 1971; Amdt. 29-11, 41 FR 5290, Feb. 5, 1976]

                            Safety Equipment



Sec. 29.1411  General.

    (a) Accessibility. Required safety equipment to be used by the crew 
in an emergency, such as automatic liferaft releases, must be readily 
accessible.
    (b) Stowage provisions. Stowage provisions for required emergency 
equipment must be furnished and must--
    (1) Be arranged so that the equipment is directly accessible and its 
location is obvious; and
    (2) Protect the safety equipment from inadvertent damage.
    (c) Emergency exit descent device. The stowage provisions for the 
emergency exit descent device required by Sec. 29.809(f) must be at the 
exits for which they are intended.
    (d) Liferafts. Liferafts must be stowed near exits through which the 
rafts can be launched during an unplanned ditching. Rafts automatically 
or remotely released outside the rotorcraft must be attached to the 
rotorcraft by the static line prescribed in Sec. 29.1415.
    (e) Long-range signaling device. The stowage provisions for the 
long-range signaling device required by Sec. 29.1415 must be near an 
exit available during an unplanned ditching.
    (f) Life preservers. Each life preserver must be within easy reach 
of each occupant while seated.



Sec. 29.1413  Safety belts: passenger warning device.

    (a) If there are means to indicate to the passengers when safety 
belts should be fastened, they must be installed to be operated from 
either pilot seat.
    (b) Each safety belt must be equipped with a metal to metal latching 
device.

(Secs. 313, 314, and 601 through 610 of the Federal Aviation Act of 1958 
(49 U.S.C. 1354, 1355, and 1421 through 1430) and sec. 6(c), Dept. of 
Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-16 43 
FR 46233, Oct. 5, 1978]



Sec. 29.1415  Ditching equipment.

    (a) Emergency flotation and signaling equipment required by any 
operating rule of this chapter must meet the requirements of this 
section.
    (b) Each liferaft and each life preserver must be approved. In 
addition--
    (1) Provide not less than two rafts, of an approximately equal rated 
capacity and buoyancy to accommodate the occupants of the rotorcraft; 
and
    (2) Each raft must have a trailing line, and must have a static line 
designed to hold the raft near the rotorcraft but to release it if the 
rotorcraft becomes totally submerged.
    (c) Approved survival equipment must be attached to each liferaft.
    (d) There must be an approved survival type emergency locator 
transmitter for use in one life raft.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-8, 36 
FR 18722, Sept. 21, 1971; Amdt. 29-19, 45 FR 38348, June 9, 1980; Amdt. 
27-26, 55 FR 8005, Mar. 6, 1990; Amdt. 29-33, 59 FR 32057, June 21, 
1994]

[[Page 838]]



Sec. 29.1419  Ice protection.

    (a) To obtain certification for flight into icing conditions, 
compliance with this section must be shown.
    (b) It must be demonstrated that the rotorcraft can be safely 
operated in the continuous maximum and intermittent maximum icing 
conditions determined under appendix C of this part within the 
rotorcraft altitude envelope. An analysis must be performed to 
establish, on the basis of the rotorcraft's operational needs, the 
adequacy of the ice protection system for the various components of the 
rotorcraft.
    (c) In addition to the analysis and physical evaluation prescribed 
in paragraph (b) of this section, the effectiveness of the ice 
protection system and its components must be shown by flight tests of 
the rotorcraft or its components in measured natural atmospheric icing 
conditions and by one or more of the following tests as found necessary 
to determine the adequacy of the ice protection system:
    (1) Laboratory dry air or simulated icing tests, or a combination of 
both, of the components or models of the components.
    (2) Flight dry air tests of the ice protection system as a whole, or 
its individual components.
    (3) Flight tests of the rotorcraft or its components in measured 
simulated icing conditions.
    (d) The ice protection provisions of this section are considered to 
be applicable primarily to the airframe. Powerplant installation 
requirements are contained in Subpart E of this part.
    (e) A means must be identified or provided for determining the 
formation of ice on critical parts of the rotorcraft. Unless otherwise 
restricted, the means must be available for nighttime as well as daytime 
operation. The rotorcraft flight manual must describe the means of 
determining ice formation and must contain information necessary for 
safe operation of the rotorcraft in icing conditions.

[Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]

                         Miscellaneous Equipment



Sec. 29.1431  Electronic equipment.

    (a) Radio communication and navigation equipment installations must 
be free from hazards in themselves, in their method of operation, and in 
their effects on other components, under any critical environmental 
conditions.
    (b) Radio communication and navigation equipment, controls, and 
wiring must be installed so that operation of any one unit or system of 
units will not adversely affect the simultaneous operation of any other 
radio or electronic unit, or system of units, required by this chapter.



Sec. 29.1433  Vacuum systems.

    (a) There must be means, in addition to the normal pressure relief, 
to automatically relieve the pressure in the discharge lines from the 
vacuum air pump when the delivery temperature of the air becomes unsafe.
    (b) Each vacuum air system line and fitting on the discharge side of 
the pump that might contain flammable vapors or fluids must meet the 
requirements of Sec. 29.1183 if they are in a designated fire zone.
    (c) Other vacuum air system components in designated fire zones must 
be at least fire resistant.



Sec. 29.1435  Hydraulic systems.

    (a) Design. Each hydraulic system must be designed as follows:
    (1) Each element of the hydraulic system must be designed to 
withstand, without detrimental, permanent deformation, any structural 
loads that may be imposed simultaneously with the maximum operating 
hydraulic loads.
    (2) Each element of the hydraulic system must be designed to 
withstand pressures sufficiently greater than those prescribed in 
paragraph (b) of this section to show that the system will not rupture 
under service conditions.
    (3) There must be means to indicate the pressure in each main 
hydraulic power system.
    (4) There must be means to ensure that no pressure in any part of 
the system will exceed a safe limit above the maximum operating pressure 
of the

[[Page 839]]

system, and to prevent excessive pressures resulting from any fluid 
volumetric change in lines likely to remain closed long enough for such 
a change to take place. The possibility of detrimental transient (surge) 
pressures during operation must be considered.
    (5) Each hydraulic line, fitting, and component must be installed 
and supported to prevent excessive vibration and to withstand inertia 
loads. Each element of the installation must be protected from abrasion, 
corrosion, and mechanical damage.
    (6) Means for providing flexibility must be used to connect points, 
in a hydraulic fluid line, between which relative motion or differential 
vibration exists.
    (b) Tests. Each element of the system must be tested to a proof 
pressure of 1.5 times the maximum pressure to which that element will be 
subjected in normal operation, without failure, malfunction, or 
detrimental deformation of any part of the system.
    (c) Fire protection. Each hydraulic system using flammable hydraulic 
fluid must meet the applicable requirements of Sec. Sec. 29.861, 
29.1183, 29.1185, and 29.1189.



Sec. 29.1439  Protective breathing equipment.

    (a) If one or more cargo or baggage compartments are to be 
accessible in flight, protective breathing equipment must be available 
for an appropriate crewmember.
    (b) For protective breathing equipment required by paragraph (a) of 
this section or by any operating rule of this chapter--
    (1) That equipment must be designed to protect the crew from smoke, 
carbon dioxide, and other harmful gases while on flight deck duty;
    (2) That equipment must include--
    (i) Masks covering the eyes, nose, and mouth; or
    (ii) Masks covering the nose and mouth, plus accessory equipment to 
protect the eyes; and
    (3) That equipment must supply protective oxygen of 10 minutes 
duration per crewmember at a pressure altitude of 8,000 feet with a 
respiratory minute volume of 30 liters per minute BTPD.



Sec. 29.1457  Cockpit voice recorders.

    (a) Each cockpit voice recorder required by the operating rules of 
this chapter must be approved, and must be installed so that it will 
record the following:
    (1) Voice communications transmitted from or received in the 
rotorcraft by radio.
    (2) Voice communications of flight crewmembers on the flight deck.
    (3) Voice communications of flight crewmembers on the flight deck, 
using the rotorcraft's interphone system.
    (4) Voice or audio signals identifying navigation or approach aids 
introduced into a headset or speaker.
    (5) Voice communications of flight crewmembers using the passenger 
loudspeaker system, if there is such a system, and if the fourth channel 
is available in accordance with the requirements of paragraph (c)(4)(ii) 
of this section.
    (6) If datalink communication equipment is installed, all datalink 
communications, using an approved data message set. Datalink messages 
must be recorded as the output signal from the communications unit that 
translates the signal into usable data.
    (b) The recording requirements of paragraph (a)(2) of this section 
may be met--
    (1) By installing a cockpit-mounted area microphone, located in the 
best position for recording voice communications originating at the 
first and second pilot stations and voice communications of other 
crewmembers on the flight deck when directed to those stations; or
    (2) By installing a continually energized or voice-actuated lip 
microphone at the first and second pilot stations.

The microphone specified in this paragraph must be so located and, if 
necessary, the preamplifiers and filters of the recorder must be so 
adjusted or supplemented, that the recorded communications are 
intelligible when recorded under flight cockpit noise conditions and 
played back. The level of intelligibility must be approved by the 
Administrator. Repeated aural or visual playback of the record may be 
used in evaluating intelligibility.
    (c) Each cockpit voice recorder must be installed so that the part 
of the

[[Page 840]]

communication or audio signals specified in paragraph (a) of this 
section obtained from each of the following sources is recorded on a 
separate channel:
    (1) For the first channel, from each microphone, headset, or speaker 
used at the first pilot station.
    (2) For the second channel, from each microphone, headset, or 
speaker used at the second pilot station.
    (3) For the third channel, from the cockpit-mounted area microphone, 
or the continually energized or voice-actuated lip microphones at the 
first and second pilot stations.
    (4) For the fourth channel, from--
    (i) Each microphone, headset, or speaker used at the stations for 
the third and fourth crewmembers; or
    (ii) If the stations specified in paragraph (c)(4)(i) of this 
section are not required or if the signal at such a station is picked up 
by another channel, each microphone on the flight deck that is used with 
the passenger loudspeaker system if its signals are not picked up by 
another channel.
    (iii) Each microphone on the flight deck that is used with the 
rotorcraft's loudspeaker system if its signals are not picked up by 
another channel.
    (d) Each cockpit voice recorder must be installed so that--
    (1)(i) It receives its electrical power from the bus that provides 
the maximum reliability for operation of the cockpit voice recorder 
without jeopardizing service to essential or emergency loads.
    (ii) It remains powered for as long as possible without jeopardizing 
emergency operation of the rotorcraft.
    (2) There is an automatic means to simultaneously stop the recorder 
and prevent each erasure feature from functioning, within 10 minutes 
after crash impact;
    (3) There is an aural or visual means for preflight checking of the 
recorder for proper operation;
    (4) Whether the cockpit voice recorder and digital flight data 
recorder are installed in separate boxes or in a combination unit, no 
single electrical failure external to the recorder may disable both the 
cockpit voice recorder and the digital flight data recorder; and
    (5) It has an independent power source--
    (i) That provides 10 1 minutes of electrical 
power to operate both the cockpit voice recorder and cockpit-mounted 
area microphone;
    (ii) That is located as close as practicable to the cockpit voice 
recorder; and
    (iii) To which the cockpit voice recorder and cockpit-mounted area 
microphone are switched automatically in the event that all other power 
to the cockpit voice recorder is interrupted either by normal shutdown 
or by any other loss of power to the electrical power bus.
    (e) The record container must be located and mounted to minimize the 
probability of rupture of the container as a result of crash impact and 
consequent heat damage to the record from fire.
    (f) If the cockpit voice recorder has a bulk erasure device, the 
installation must be designed to minimize the probability of inadvertent 
operation and actuation of the device during crash impact.
    (g) Each recorder container must be either bright orange or bright 
yellow.
    (h) When both a cockpit voice recorder and a flight data recorder 
are required by the operating rules, one combination unit may be 
installed, provided that all other requirements of this section and the 
requirements for flight data recorders under this part are met.

[Amdt. 29-6, 35 FR 7293, May 9, 1970, as amended by Amdt. 29-50, 73 FR 
12564, Mar. 7, 2008; 74 FR 32800, July 9, 2009; Amdt. 29-52, 75 FR 
17045, Apr. 5, 2010]



Sec. 29.1459  Flight data recorders.

    (a) Each flight recorder required by the operating rules of 
Subchapter G of this chapter must be installed so that:
    (1) It is supplied with airspeed, altitude, and directional data 
obtained from sources that meet the accuracy requirements of Sec. Sec. 
29.1323, 29.1325, and 29.1327 of this part, as applicable;
    (2) The vertical acceleration sensor is rigidly attached, and 
located longitudinally within the approved center of gravity limits of 
the rotorcraft;

[[Page 841]]

    (3)(i) It receives its electrical power from the bus that provides 
the maximum reliability for operation of the flight data recorder 
without jeopardizing service to essential or emergency loads.
    (ii) It remains powered for as long as possible without jeopardizing 
emergency operation of the rotorcraft.
    (4) There is an aural or visual means for perflight checking of the 
recorder for proper recording of data in the storage medium;
    (5) Except for recorders powered solely by the engine-drive 
electrical generator system, there is an automatic means to 
simultaneously stop a recorder that has a data erasure feature and 
prevent each erasure feature from functioning, within 10 minutes after 
any crash impact; and
    (6) Whether the cockpit voice recorder and digital flight data 
recorder are installed in separate boxes or in a combination unit, no 
single electrical failure external to the recorder may disable both the 
cockpit voice recorder and the digital flight data recorder.
    (b) Each nonejectable recorder container must be located and mounted 
so as to minimize the probability of container rupture resulting from 
crash impact and subsequent damage to the record from fire.
    (c) A correlation must be established between the flight recorder 
readings of airspeed, altitude, and heading and the corresponding 
readings (taking into account correction factors) of the first pilot's 
instruments. This correlation must cover the airspeed range over which 
the aircraft is to be operated, the range of altitude to which the 
aircraft is limited, and 360 degrees of heading. Correlation may be 
established on the ground as appropriate.
    (d) Each recorder container must:
    (1) Be either bright orange or bright yellow;
    (2) Have a reflective tape affixed to its external surface to 
facilitate its location under water; and
    (3) Have an underwater locating device, when required by the 
operating rules of this chapter, on or adjacent to the container which 
is secured in such a manner that it is not likely to be separated during 
crash impact.
    (e) When both a cockpit voice recorder and a flight data recorder 
are required by the operating rules, one combination unit may be 
installed, provided that all other requirements of this section and the 
requirements for cockpit voice recorders under this part are met.

[Amdt. 29-25, 53 FR 26145, July 11, 1988; 53 FR 26144, July 11, 1988, as 
amended by Amdt. 29-50, 73 FR 12564, Mar. 7, 2008; 74 FR 32800, July 9, 
2009; Amdt. 29-52, 75 FR 17045, Apr. 5, 2010]



Sec. 29.1461  Equipment containing high energy rotors.

    (a) Equipment containing high energy rotors must meet paragraph (b), 
(c), or (d) of this section.
    (b) High energy rotors contained in equipment must be able to 
withstand damage caused by malfunctions, vibration, abnormal speeds, and 
abnormal temperatures. In addition--
    (1) Auxiliary rotor cases must be able to contain damage caused by 
the failure of high energy rotor blades; and
    (2) Equipment control devices, systems, and instrumentation must 
reasonably ensure that no operating limitations affecting the integrity 
of high energy rotors will be exceeded in service.
    (c) It must be shown by test that equipment containing high energy 
rotors can contain any failure of a high energy rotor that occurs at the 
highest speed obtainable with the normal speed control devices 
inoperative.
    (d) Equipment containing high energy rotors must be located where 
rotor failure will neither endanger the occupants nor adversely affect 
continued safe flight.

[Amdt. 29-3, 33 FR 971, Jan. 26, 1968]



             Subpart G_Operating Limitations and Information



Sec. 29.1501  General.

    (a) Each operating limitation specified in Sec. Sec. 29.1503 
through 29.1525 and other limitations and information necessary for safe 
operation must be established.
    (b) The operating limitations and other information necessary for 
safe operation must be made available to

[[Page 842]]

the crewmembers as prescribed in Sec. Sec. 29.1541 through 29.1589.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 29-15, 43 FR 2327, Jan. 16, 1978]

                          Operating Limitations



Sec. 29.1503  Airspeed limitations: general.

    (a) An operating speed range must be established.
    (b) When airspeed limitations are a function of weight, weight 
distribution, altitude, rotor speed, power, or other factors, airspeed 
limitations corresponding with the critical combinations of these 
factors must be established.



Sec. 29.1505  Never-exceed speed.

    (a) The never-exceed speed, VNE, must be established so that it is--
    (1) Not less than 40 knots (CAS); and
    (2) Not more than the lesser of--
    (i) 0.9 times the maximum forward speeds established under Sec. 
29.309;
    (ii) 0.9 times the maximum speed shown under Sec. Sec. 29.251 and 
29.629; or
    (iii) 0.9 times the maximum speed substantiated for advancing blade 
tip mach number effects under critical altitude conditions.
    (b) VNE may vary with altitude, r.p.m., temperature, and weight, 
if--
    (1) No more than two of these variables (or no more than two 
instruments integrating more than one of these variables) are used at 
one time; and
    (2) The ranges of these variables (or of the indications on 
instruments integrating more than one of these variables) are large 
enough to allow an operationally practical and safe variation of VNE.
    (c) For helicopters, a stabilized power-off VNE denoted as VNE 
(power-off) may be established at a speed less than VNE established 
pursuant to paragraph (a) of this section, if the following conditions 
are met:
    (1) VNE (power-off) is not less than a speed midway between the 
power-on VNE and the speed used in meeting the requirements of--
    (i) Sec. 29.67(a)(3) for Category A helicopters;
    (ii) Sec. 29.65(a) for Category B helicopters, except multi-engine 
helicopters meeting the requirements of Sec. 29.67(b); and
    (iii) Sec. 29.67(b) for multi-engine Category B helicopters meeting 
the requirements of Sec. 29.67(b).
    (2) VNE (power-off) is--
    (i) A constant airspeed;
    (ii) A constant amount less than power-on VNE; or
    (iii) A constant airspeed for a portion of the altitude range for 
which certification is requested, and a constant amount less than power-
on VNE for the remainder of the altitude range.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 29-3, 33 FR 971, Jan. 26, 1968, as amended by Amdt. 29-15, 43 FR 
2327, Jan. 16, 1978; Amdt. 29-24, 49 FR 44440, Nov. 6, 1984]



Sec. 29.1509  Rotor speed.

    (a) Maximum power-off (autorotation). The maximum power-off rotor 
speed must be established so that it does not exceed 95 percent of the 
lesser of--
    (1) The maximum design r.p.m. determined under Sec. 29.309(b); and
    (2) The maximum r.p.m. shown during the type tests.
    (b) Minimum power-off. The minimum power-off rotor speed must be 
established so that it is not less than 105 percent of the greater of--
    (1) The minimum shown during the type tests; and
    (2) The minimum determined by design substantiation.
    (c) Minimum power-on. The minimum power-on rotor speed must be 
established so that it is--
    (1) Not less than the greater of--
    (i) The minimum shown during the type tests; and
    (ii) The minimum determined by design substantiation; and
    (2) Not more than a value determined under Sec. 29.33 (a)(1) and 
(c)(1).



Sec. 29.1517  Limiting height-speed envelope.

    For Category A rotorcraft, if a range of heights exists at any 
speed, including zero, within which it is not possible to make a safe 
landing following power failure, the range of heights and its

[[Page 843]]

variation with forward speed must be established, together with any 
other pertinent information, such as the kind of landing surface.

[Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]



Sec. 29.1519  Weight and center of gravity.

    The weight and center of gravity limitations determined under 
Sec. Sec. 29.25 and 29.27, respectively, must be established as 
operating limitations.



Sec. 29.1521  Powerplant limitations.

    (a) General. The powerplant limitations prescribed in this section 
must be established so that they do not exceed the corresponding limits 
for which the engines are type certificated.
    (b) Takeoff operation. The powerplant takeoff operation must be 
limited by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value shown during the type tests;
    (2) The maximum allowable manifold pressure (for reciprocating 
engines);
    (3) The maximum allowable turbine inlet or turbine outlet gas 
temperature (for turbine engines);
    (4) The maximum allowable power or torque for each engine, 
considering the power input limitations of the transmission with all 
engines operating;
    (5) The maximum allowable power or torque for each engine 
considering the power input limitations of the transmission with one 
engine inoperative;
    (6) The time limit for the use of the power corresponding to the 
limitations established in paragraphs (b)(1) through (5) of this 
section; and
    (7) If the time limit established in paragraph (b)(6) of this 
section exceeds 2 minutes--
    (i) The maximum allowable cylinder head or coolant outlet 
temperature (for reciprocating engines); and
    (ii) The maximum allowable engine and transmission oil temperatures.
    (c) Continuous operation. The continuous operation must be limited 
by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value shown during the type tests;
    (2) The minimum rotational speed shown under the rotor speed 
requirements in Sec. 29.1509(c).
    (3) The maximum allowable manifold pressure (for reciprocating 
engines);
    (4) The maximum allowable turbine inlet or turbine outlet gas 
temperature (for turbine engines);
    (5) The maximum allowable power or torque for each engine, 
considering the power input limitations of the transmission with all 
engines operating;
    (6) The maximum allowable power or torque for each engine, 
considering the power input limitations of the transmission with one 
engine inoperative; and
    (7) The maximum allowable temperatures for--
    (i) The cylinder head or coolant outlet (for reciprocating engines);
    (ii) The engine oil; and
    (iii) The transmission oil.
    (d) Fuel grade or designation. The minimum fuel grade (for 
reciprocating engines) or fuel designation (for turbine engines) must be 
established so that it is not less than that required for the operation 
of the engines within the limitations in paragraphs (b) and (c) of this 
section.
    (e) Ambient temperature. Ambient temperature limitations (including 
limitations for winterization installations if applicable) must be 
established as the maximum ambient atmospheric temperature at which 
compliance with the cooling provisions of Sec. Sec. 29.1041 through 
29.1049 is shown.
    (f) Two and one-half minute OEI power operation. Unless otherwise 
authorized, the use of 2\1/2\-minute OEI power must be limited to engine 
failure operation of multiengine, turbine-powered rotorcraft for not 
longer than 2\1/2\ minutes for any period in which that power is used. 
The use of 2\1/2\-minute OEI power must also be limited by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value shown during the type tests;
    (2) The maximum allowable gas temperature;
    (3) The maximum allowable torque; and

[[Page 844]]

    (4) The maximum allowable oil temperature.
    (g) Thirty-minute OEI power operation. Unless otherwise authorized, 
the use of 30-minute OEI power must be limited to multiengine, turbine-
powered rotorcraft for not longer than 30 minutes after failure of an 
engine. The use of 30-minute OEI power must also be limited by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value shown during the type tests;
    (2) The maximum allowable gas temperature;
    (3) The maximum allowable torque; and
    (4) The maximum allowable oil temperature.
    (h) Continuous OEI power operation. Unless otherwise authorized, the 
use of continuous OEI power must be limited to multiengine, turbine-
powered rotorcraft for continued flight after failure of an engine. The 
use of continuous OEI power must also be limited by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value shown during the type tests.
    (2) The maximum allowable gas temperature;
    (3) The maximum allowable torque; and
    (4) The maximum allowable oil temperature.
    (i) Rated 30-second OEI power operation. Rated 30-second OEI power 
is permitted only on multiengine, turbine-powered rotorcraft, also 
certificated for the use of rated 2-minute OEI power, and can only be 
used for continued operation of the remaining engine(s) after a failure 
or precautionary shutdown of an engine. It must be shown that following 
application of 30-second OEI power, any damage will be readily 
detectable by the applicable inspections and other related procedures 
furnished in accordance with Section A29.4 of appendix A of this part 
and Section A33.4 of appendix A of part 33. The use of 30-second OEI 
power must be limited to not more than 30 seconds for any period in 
which that power is used, and by--
    (1) The maximum rotational speed which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value demonstrated during the type tests;
    (2) The maximum allowable gas temperature; and
    (3) The maximum allowable torque.
    (j) Rated 2-minute OEI power operation. Rated 2-minute OEI power is 
permitted only on multiengine, turbine-powered rotorcraft, also 
certificated for the use of rated 30-second OEI power, and can only be 
used for continued operation of the remaining engine(s) after a failure 
or precautionary shutdown of an engine. It must be shown that following 
application of 2-minute OEI power, any damage will be readily detectable 
by the applicable inspections and other related procedures furnished in 
accordance with Section A29.4 of appendix a of this part and Section 
A33.4 of appendix A of part 33. The use of 2-minute OEI power must be 
limited to not more than 2 minutes for any period in which that power is 
used, and by--
    (1) The maximum rotational speed, which may not be greater than--
    (i) The maximum value determined by the rotor design; or
    (ii) The maximum value demonstrated during the type tests;
    (2) The maximum allowable gas temperature; and
    (3) The maximum allowable torque.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-1, 30 
FR 8778, July 13, 1965; Amdt. 29-3, 33 FR 971, Jan. 26, 1968; Amdt. 29-
15, 43 FR 2327, Jan. 16, 1978; Amdt. 29-26, 53 FR 34220, Sept. 2, 1988; 
Amdt. 29-34, 59 FR 47768, Sept. 16, 1994; Amdt. 29-41, 62 FR 46173, Aug. 
29, 1997]



Sec. 29.1522  Auxiliary power unit limitations.

    If an auxiliary power unit that meets the requirements of TSO-C77 is 
installed in the rotorcraft, the limitations established for that 
auxiliary power unit under the TSO including the categories of operation 
must be

[[Page 845]]

specified as operating limitations for the rotorcraft.

(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, 1423), sec. 6(c), Dept. of Transportation Act (49 U.S.C. 
1655(c)))

[Amdt. 29-17, 43 FR 50602, Oct. 30, 1978]



Sec. 29.1523  Minimum flight crew.

    The minimum flight crew must be established so that it is sufficient 
for safe operation, considering--
    (a) The workload on individual crewmembers;
    (b) The accessibility and ease of operation of necessary controls by 
the appropriate crewmember; and
    (c) The kinds of operation authorized under Sec. 29.1525.



Sec. 29.1525  Kinds of operations.

    The kinds of operations (such as VFR, IFR, day, night, or icing) for 
which the rotorcraft is approved are established by demonstrated 
compliance with the applicable certification requirements and by the 
installed equipment.

[Amdt. 29-24, 49 FR 44440, Nov. 6, 1984]



Sec. 29.1527  Maximum operating altitude.

    The maximum altitude up to which operation is allowed, as limited by 
flight, structural, powerplant, functional, or equipment 
characteristics, must be established.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 29-15, 43 FR 2327, Jan. 16, 1978]



Sec. 29.1529  Instructions for Continued Airworthiness.

    The applicant must prepare Instructions for Continued Airworthiness 
in accordance with appendix A to this part that are acceptable to the 
Administrator. The instructions may be incomplete at type certification 
if a program exists to ensure their completion prior to delivery of the 
first rotorcraft or issuance of a standard certificate of airworthiness, 
whichever occurs later.

[Amdt. 29-20, 45 FR 60178, Sept. 11, 1980]

                          Markings and Placards



Sec. 29.1541  General.

    (a) The rotorcraft must contain--
    (1) The markings and placards specified in Sec. Sec. 29.1545 
through 29.1565; and
    (2) Any additional information, instrument markings, and placards 
required for the safe operation of the rotorcraft if it has unusual 
design, operating or handling characteristics.
    (b) Each marking and placard prescribed in paragraph (a) of this 
section--
    (1) Must be displayed in a conspicuous place; and
    (2) May not be easily erased, disfigured, or obscured.



Sec. 29.1543  Instrument markings: general.

    For each instrument--
    (a) When markings are on the cover glass of the instrument there 
must be means to maintain the correct alignment of the glass cover with 
the face of the dial; and
    (b) Each arc and line must be wide enough, and located to be clearly 
visible to the pilot.



Sec. 29.1545  Airspeed indicator.

    (a) Each airspeed indicator must be marked as specified in paragraph 
(b) of this section, with the marks located at the corresponding 
indicated airspeeds.
    (b) The following markings must be made:
    (1) A red radial line--
    (i) For rotorcraft other than helicopters, at VNE; and
    (ii) For helicopters, at a VNE (power-on).
    (2) A red, cross-hatched radial line at VNE (power-off) for 
helicopters, if VNE (power-off) is less than VNE (power-on).
    (3) For the caution range, a yellow arc.
    (4) For the safe operating range, a green arc.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-15, 43 
FR 2327, Jan. 16, 1978; 43 FR 3900, Jan. 30, 1978; Amdt. 29-17, 43 FR 
50602, Oct. 30, 1978]

[[Page 846]]



Sec. 29.1547  Magnetic direction indicator.

    (a) A placard meeting the requirements of this section must be 
installed on or near the magnetic direction indicator.
    (b) The placard must show the calibration of the instrument in level 
flight with the engines operating.
    (c) The placard must state whether the calibration was made with 
radio receivers on or off.
    (d) Each calibration reading must be in terms of magnetic heading in 
not more than 45 degree increments.



Sec. 29.1549  Powerplant instruments.

    For each required powerplant instrument, as appropriate to the type 
of instruments--
    (a) Each maximum and, if applicable, minimum safe operating limit 
must be marked with a red radial or a red line;
    (b) Each normal operating range must be marked with a green arc or 
green line, not extending beyond the maximum and minimum safe limits;
    (c) Each takeoff and precautionary range must be marked with a 
yellow arc or yellow line;
    (d) Each engine or propeller range that is restricted because of 
excessive vibration stresses must be marked with red arcs or red lines; 
and
    (e) Each OEI limit or approved operating range must be marked to be 
clearly differentiated from the markings of paragraphs (a) through (d) 
of this section except that no marking is normally required for the 30-
second OEI limit.

[Amdt. 29-12, 41 FR 55474, Dec. 20, 1976, as amended by Amdt. 29-26, 53 
FR 34220, Sept. 2, 1988; Amdt. 29-34, 59 FR 47769, Sept. 16, 1994]



Sec. 29.1551  Oil quantity indicator.

    Each oil quantity indicator must be marked with enough increments to 
indicate readily and accurately the quantity of oil.



Sec. 29.1553  Fuel quantity indicator.

    If the unusable fuel supply for any tank exceeds one gallon, or five 
percent of the tank capacity, whichever is greater, a red arc must be 
marked on its indicator extending from the calibrated zero reading to 
the lowest reading obtainable in level flight.



Sec. 29.1555  Control markings.

    (a) Each cockpit control, other than primary flight controls or 
control whose function is obvious, must be plainly marked as to its 
function and method of operation.
    (b) For powerplant fuel controls--
    (1) Each fuel tank selector valve control must be marked to indicate 
the position corresponding to each tank and to each existing cross feed 
position;
    (2) If safe operation requires the use of any tanks in a specific 
sequence, that sequence must be marked on, or adjacent to, the selector 
for those tanks; and
    (3) Each valve control for any engine of a multiengine rotorcraft 
must be marked to indicate the position corresponding to each engine 
controlled.
    (c) Usable fuel capacity must be marked as follows:
    (1) For fuel systems having no selector controls, the usable fuel 
capacity of the system must be indicated at the fuel quantity indicator.
    (2) For fuel systems having selector controls, the usable fuel 
capacity available at each selector control position must be indicated 
near the selector control.
    (d) For accessory, auxiliary, and emergency controls--
    (1) Each essential visual position indicator, such as those showing 
rotor pitch or landing gear position, must be marked so that each 
crewmember can determine at any time the position of the unit to which 
it relates; and
    (2) Each emergency control must be red and must be marked as to 
method of operation.
    (e) For rotorcraft incorporating retractable landing gear, the 
maximum landing gear operating speed must be displayed in clear view of 
the pilot.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 
FR 55474, Dec. 20, 1976; Amdt. 29-24, 49 FR 44440, Nov. 6, 1984]



Sec. 29.1557  Miscellaneous markings and placards.

    (a) Baggage and cargo compartments, and ballast location. Each 
baggage and cargo compartment, and each ballast location must have a 
placard stating any limitations on contents, including weight, that are 
necessary under the loading requirements.

[[Page 847]]

    (b) Seats. If the maximum allowable weight to be carried in a seat 
is less than 170 pounds, a placard stating the lesser weight must be 
permanently attached to the seat structure.
    (c) Fuel and oil filler openings. The following apply:
    (1) Fuel filler openings must be marked at or near the filler cover 
with--
    (i) The word ``fuel'';
    (ii) For reciprocating engine powered rotorcraft, the minimum fuel 
grade;
    (iii) For turbine-engine-powered rotorcraft, the permissible fuel 
designations, except that if impractical, this information may be 
included in the rotorcraft flight manual, and the fuel filler may be 
marked with an appropriate reference to the flight manual; and
    (iv) For pressure fueling systems, the maximum permissible fueling 
supply pressure and the maximum permissible defueling pressure.
    (2) Oil filler openings must be marked at or near the filler cover 
with the word ``oil''.
    (d) Emergency exit placards. Each placard and operating control for 
each emergency exit must differ in color from the surrounding fuselage 
surface as prescribed in Sec. 29.811(h)(2). A placard must be near each 
emergency exit control and must clearly indicate the location of that 
exit and its method of operation.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 971, Jan. 26, 1968; Amdt. 29-12, 41 FR 55474, Dec. 20, 1976; Amdt. 
29-26, 53 FR 34220, Sept. 2, 1988]



Sec. 29.1559  Limitations placard.

    There must be a placard in clear view of the pilot that specifies 
the kinds of operations (VFR, IFR, day, night, or icing) for which the 
rotorcraft is approved.

[Amdt. 29-24, 49 FR 44440, Nov. 6, 1984]



Sec. 29.1561  Safety equipment.

    (a) Each safety equipment control to be operated by the crew in 
emergency, such as controls for automatic liferaft releases, must be 
plainly marked as to its method of operation.
    (b) Each location, such as a locker or compartment, that carries any 
fire extinguishing, signaling, or other life saving equipment, must be 
so marked.
    (c) Stowage provisions for required emergency equipment must be 
conspicuously marked to identify the contents and facilitate removal of 
the equipment.
    (d) Each liferaft must have obviously marked operating instructions.
    (e) Approved survival equipment must be marked for identification 
and method of operation.



Sec. 29.1565  Tail rotor.

    Each tail rotor must be marked so that its disc is conspicuous under 
normal daylight ground conditions.

[Amdt. 29-3, 33 FR 971, Jan. 26, 1968]

                        Rotorcraft Flight Manual



Sec. 29.1581  General.

    (a) Furnishing information. A Rotorcraft Flight Manual must be 
furnished with each rotorcraft, and it must contain the following:
    (1) Information required by Sec. Sec. 29.1583 through 29.1589.
    (2) Other information that is necessary for safe operation because 
of design, operating, or handling characteristics.
    (b) Approved information. Each part of the manual listed in 
Sec. Sec. 29.1583 through 29.1589 that is appropriate to the 
rotorcraft, must be furnished, verified, and approved, and must be 
segregated, indentified, and clearly distinguished from each unapproved 
part of that manual.
    (c) [Reserved]
    (d) Table of contents. Each Rotorcraft Flight Manual must include a 
table of contents if the complexity of the manual indicates a need for 
it.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 29-15, 43 FR 2327, Jan. 16, 1978]



Sec. 29.1583  Operating limitations.

    (a) Airspeed and rotor limitations. Information necessary for the 
marking of airspeed and rotor limitations on or near their respective 
indicators must be furnished. The significance of each

[[Page 848]]

limitation and of the color coding must be explained.
    (b) Powerplant limitations. The following information must be 
furnished:
    (1) Limitations required by Sec. 29.1521.
    (2) Explanation of the limitations, when appropriate.
    (3) Information necessary for marking the instruments required by 
Sec. Sec. 29.1549 through 29.1553.
    (c) Weight and loading distribution. The weight and center of 
gravity limits required by Sec. Sec. 29.25 and 29.27, respectively, 
must be furnished. If the variety of possible loading conditions 
warrants, instructions must be included to allow ready observance of the 
limitations.
    (d) Flight crew. When a flight crew of more than one is required, 
the number and functions of the minimum flight crew determined under 
Sec. 29.1523 must be furnished.
    (e) Kinds of operation. Each kind of operation for which the 
rotorcraft and its equipment installations are approved must be listed.
    (f) Limiting heights. Enough information must be furnished to allow 
compliance with Sec. 29.1517.
    (g) Maximum allowable wind. For Category A rotorcraft, the maximum 
allowable wind for safe operation near the ground must be furnished.
    (h) Altitude. The altitude established under Sec. 29.1527 and an 
explanation of the limiting factors must be furnished.
    (i) Ambient temperature. Maximum and minimum ambient temperature 
limitations must be furnished.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 
FR 971, Jan. 26, 1968; Amdt. 29-15, 43 FR 2327, Jan. 16, 1978; Amdt. 29-
17, 43 FR 50602, Oct. 30, 1978; Amdt. 29-24, 49 FR 44440, Nov. 6, 1984]



Sec. 29.1585  Operating procedures.

    (a) The parts of the manual containing operating procedures must 
have information concerning any normal and emergency procedures, and 
other information necessary for safe operation, including the applicable 
procedures, such as those involving minimum speeds, to be followed if an 
engine fails.
    (b) For multiengine rotorcraft, information identifying each 
operating condition in which the fuel system independence prescribed in 
Sec. 29.953 is necessary for safety must be furnished, together with 
instructions for placing the fuel system in a configuration used to show 
compliance with that section.
    (c) For helicopters for which a VNE (power-off) is established under 
Sec. 29.1505(c), information must be furnished to explain the VNE 
(power-off) and the procedures for reducing airspeed to not more than 
the VNE (power-off) following failure of all engines.
    (d) For each rotorcraft showing compliance with Sec. 29.1353 
(c)(6)(ii) or (c)(6)(iii), the operating procedures for disconnecting 
the battery from its charging source must be furnished.
    (e) If the unusable fuel supply in any tank exceeds 5 percent of the 
tank capacity, or 1 gallon, whichever is greater, information must be 
furnished which indicates that when the fuel quantity indicator reads 
``zero'' in level flight, any fuel remaining in the fuel tank cannot be 
used safely in flight.
    (f) Information on the total quantity of usable fuel for each fuel 
tank must be furnished.
    (g) For Category B rotorcraft, the airspeeds and corresponding rotor 
speeds for minimum rate of descent and best glide angle as prescribed in 
Sec. 29.71 must be provided.

(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 
1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), 
Dept. of Transportation Act (49 U.S.C. 1655(c)))

[Amdt. 29-2, 32 FR 6914, May 5, 1967, as amended by Amdt. 29-15, 43 FR 
2328, Jan. 16, 1978; Amdt. 29-17, 43 FR 50602, Oct. 30, 1978; Amdt. 29-
24, 49 FR 44440, Nov. 6, 1984]



Sec. 29.1587  Performance information.

    Flight manual performance information which exceeds any operating 
limitation may be shown only to the extent necessary for presentation 
clarity or to determine the effects of approved optional equipment or 
procedures. When data beyond operating limits are shown, the limits must 
be clearly indicated. The following must be provided:

[[Page 849]]

    (a) Category A. For each category A rotorcraft, the Rotorcraft 
Flight Manual must contain a summary of the performance data, including 
data necessary for the application of any operating rule of this 
chapter, together with descriptions of the conditions, such as 
airspeeds, under which this data was determined, and must contain--
    (1) The indicated airspeeds corresponding with those determined for 
takeoff, and the procedures to be followed if the critical engine fails 
during takeoff;
    (2) The airspeed calibrations;
    (3) The techniques, associated airspeeds, and rates of descent for 
autorotative landings;
    (4) The rejected takeoff distance determined under Sec. 29.62 and 
the takeoff distance determined under Sec. 29.61;
    (5) The landing data determined under Sec. 29.81 and Sec. 29.85;
    (6) The steady gradient of climb for each weight, altitude, and 
temperature for which takeoff data are to be scheduled, along the 
takeoff path determined in the flight conditions required in Sec. 
29.67(a)(1) and (a)(2):
    (i) In the flight conditions required in Sec. 29.67(a)(1) between 
the end of the takeoff distance and the point at which the rotorcraft is 
200 feet above the takeoff surface (or 200 feet above the lowest point 
of the takeoff profile for elevated heliports);
    (ii) In the flight conditions required in Sec. 29.67(a)(2) between 
the points at which the rotorcraft is 200 and 1000 feet above the 
takeoff surface (or 200 and 1000 feet above the lowest point of the 
takeoff profile for elevated heliports); and
    (7) Out-of-ground effect hover performance determined under Sec. 
29.49 and the maximum weight for each altitude and temperature condition 
at which the rotorcraft can safely hover out-of-ground effect in winds 
of not less than 17 knots from all azimuths. These data must be clearly 
referenced to the appropriate hover charts.
    (b) Category B. For each category B rotorcraft, the Rotorcraft 
Flight Manual must contain--
    (1) The takeoff distance and the climbout speed together with the 
pertinent information defining the flight path with respect to 
autorotative landing if an engine fails, including the calculated 
effects of altitude and temperature;
    (2) The steady rates of climb and in-ground-effect hovering ceiling, 
together with the corresponding airspeeds and other pertinent 
information, including the calculated effects of altitude and 
temperature;
    (3) The landing distance, appropriate airspeed, and type of landing 
surface, together with all pertinent information that might affect this 
distance, including the effects of weight, altitude, and temperature;
    (4) The maximum safe wind for operation near the ground;
    (5) The airspeed calibrations;
    (6) The height-speed envelope except for rotorcraft incorporating 
this as an operating limitation;
    (7) Glide distance as a function of altitude when autorotating at 
the speeds and conditions for minimum rate of descent and best glide 
angle, as determined in Sec. 29.71;
    (8) Out-of-ground effect hover performance determined under Sec. 
29.49 and the maximum safe wind demonstrated under the ambient 
conditions for data presented. In addition, the maximum weight for each 
altitude and temperature condition at which the rotorcraft can safely 
hover out-of-ground-effect in winds of not less than 17 knots from all 
azimuths. These data must be clearly referenced to the appropriate hover 
charts; and
    (9) Any additional performance data necessary for the application of 
any operating rule in this chapter.

[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-21, 48 
FR 4392, Jan. 31, 1983; Amdt. 29-24, 49 FR 44440, Nov. 6, 1984; Amdt. 
29-39, 61 FR 21901, May 10, 1996; Amdt. 29-40, 61 FR 21908, May 10, 
1996; Amdt. 29-44, 64 FR 45338, Aug. 19, 1999; Amdt. 29-51, 73 FR 11001, 
Feb. 29, 2008]



Sec. 29.1589  Loading information.

    There must be loading instructions for each possible loading 
condition between the maximum and minimum weights determined under Sec. 
29.25 that can result in a center of gravity beyond any extreme 
prescribed in Sec. 29.27, assuming any probable occupant weights.

[[Page 850]]



  Sec. Appendix A to Part 29--Instructions for Continued Airworthiness

a29.1 General

    (a) This appendix specifies requirements for the preparation of 
Instructions for Continued Airworthiness as required by Sec. 29.1529.
    (b) The Instructions for Continued Airworthiness for each rotorcraft 
must include the Instructions for Continued Airworthiness for each 
engine and rotor (hereinafter designated ``products''), for each 
applicance required by this chapter, and any required information 
relating to the interface of those appliances and products with the 
rotorcraft. If Instructions for Continued Airworthiness are not supplied 
by the manufacturer of an appliance or product installed in the 
rotorcraft, the Instructions for Continued Airworthiness for the 
rotorcraft must include the information essential to the continued 
airworthiness of the rotorcraft.
    (c) The applicant must submit to the FAA a program to show how 
changes to the Instructions for Continued Airworthiness made by the 
applicant or by the manufacturers of products and appliances installed 
in the rotorcraft will be distributed.

a29.2 Format

    (a) The Instructions for Continued Airworthiness must be in the form 
of a manual or manuals as appropriate for the quantity of data to be 
provided.
    (b) The format of the manual or manuals must provide for a practical 
arrangement.

a29.3 Content

    The contents of the manual or manuals must be prepared in the 
English language. The Instructions for Continued Airworthiness must 
contain the following manuals or sections, as appropriate, and 
information:
    (a) Rotorcraft maintenance manual or section. (1) Introduction 
information that includes an explanation of the rotorcraft's features 
and data to the extent necessary for maintenance or preventive 
maintenance.
    (2) A description of the rotorcraft and its systems and 
installations including its engines, rotors, and appliances.
    (3) Basic control and operation information describing how the 
rotorcraft components and systems are controlled and how they operate, 
including any special procedures and limitations that apply.
    (4) Servicing information that covers details regarding servicing 
points, capacities of tanks, reservoirs, types of fluids to be used, 
pressures applicable to the various systems, location of access panels 
for inspection and servicing, locations of lubrication points, the 
lubricants to be used, equipment required for servicing, tow 
instructions and limitations, mooring, jacking, and leveling 
information.
    (b) Maintenance Instructions. (1) Scheduling information for each 
part of the rotorcraft and its engines, auxiliary power units, rotors, 
accessories, instruments, and equipment that provides the recommended 
periods at which they should be cleaned, inspected, adjusted, tested, 
and lubricated, and the degree of inspection, the applicable wear 
tolerances, and work recommended at these periods. However, the 
applicant may refer to an accessory, instrument, or equipment 
manufacturer as the source of this information if the applicant shows 
that the item has an exceptionally high degree of complexity requiring 
specialized maintenance techniques, test equipment, or expertise. The 
recommended overhaul periods and necessary cross references to the 
Airworthiness Limitations section of the manual must also be included. 
In addition, the applicant must include an inspection program that 
includes the frequency and extent of the inspections necessary to 
provide for the continued airworthiness of the rotorcraft.
    (2) Troubleshooting information describing probable malfunctions, 
how to recognize those malfunctions, and the remedial action for those 
malfunctions.
    (3) Information describing the order and method of removing and 
replacing products and parts with any necessary precautions to be taken.
    (4) Other general procedural instructions including procedures for 
system testing during ground running, symmetry checks, weighing and 
determining the center of gravity, lifting and shoring, and storage 
limitations.
    (c) Diagrams of structural access plates and information needed to 
gain access for inspections when access plates are not provided.
    (d) Details for the application of special inspection techniques 
including radiographic and ultrasonic testing where such processes are 
specified.
    (e) Information needed to apply protective treatments to the 
structure after inspection.
    (f) All data relative to structural fasteners such as 
identification, discard recommendations, and torque values.
    (g) A list of special tools needed.

a29.4 Airworthiness Limitations Section

    The Instructions for Continued Airworthiness must contain a section 
titled Airworthiness Limitations that is segregated and clearly 
distinguishable from the rest of the document. This section must set 
forth each mandatory replacement time, structural inspection interval, 
and related structural inspection procedure required for type 
certification. If the Instructions for Continued Airworthiness consist 
of multiple documents, the section required by this paragraph must be 
included in the principal manual. This section must contain a legible 
statement in a prominent location that reads: ``The Airworthiness 
Limitations section is FAA approved and specifies maintenance required

[[Page 851]]

under Sec. Sec. 43.16 and 91.403 of the Federal Aviation Regulations 
unless an alternative program has been FAA approved.''

[Amdt. 29-20, 45 FR 60178, Sept. 11, 1980, as amended by Amdt. 29-27, 54 
FR 34330, Aug. 18, 1989; Amdt. 29-54, 76 FR 74664, Dec. 1, 2011]



   Sec. Appendix B to Part 29--Airworthiness Criteria for Helicopter 
                            Instrument Flight

    I. General. A transport category helicopter may not be type 
certificated for operation under the instrument flight rules (IFR) of 
this chapter unless it meets the design and installation requirements 
contained in this appendix.
    II. Definitions. (a) VYI means instrument climb speed, 
utilized instead of VY for compliance with the climb 
requirements for instrument flight.
    (b) VNEI means instrument flight never exceed speed, 
utilized instead of VNE for compliance with maximum limit 
speed requirements for instrument flight.
    (c) VMINI means instrument flight minimum speed, utilized 
in complying with minimum limit speed requirements for instrument 
flight.
    III. Trim. It must be possible to trim the cyclic, collective, and 
directional control forces to zero at all approved IFR airspeeds, power 
settings, and configurations appropriate to the type.
    IV. Static longitudinal stability. (a) General. The helicopter must 
possess positive static longitudinal control force stability at critical 
combinations of weight and center of gravity at the conditions specified 
in paragraphs IV (b) through (f) of this appendix. The stick force must 
vary with speed so that any substantial speed change results in a stick 
force clearly perceptible to the pilot. The airspeed must return to 
within 10 percent of the trim speed when the control force is slowly 
released for each trim condition specified in paragraphs IV (b) through 
(f) of this appendix.
    (b) Climb. Stability must be shown in climb thoughout the speed 
range 20 knots either side of trim with--
    (1) The helicopter trimmed at VYI;
    (2) Landing gear retracted (if retractable); and
    (3) Power required for limit climb rate (at least 1,000 fpm) at 
VYI or maximum continuous power, whichever is less.
    (c) Cruise. Stability must be shown throughout the speed range from 
0.7 to 1.1 VH or VNEI, whichever is lower, not to 
exceed 20 knots from trim with--
    (1) The helicopter trimmed and power adjusted for level flight at 
0.9 VH or 0.9 VNEI, whichever is lower; and
    (2) Landing gear retracted (if retractable).
    (d) Slow cruise. Stability must be shown throughout the speed range 
from 0.9 VMINI to 1.3 VMINI or 20 knots above trim 
speed, whichever is greater, with--
    (1) The helicopter trimmed and power adjusted for level flight at 
1.1 VMINI; and
    (2) Landing gear retracted (if retractable).
    (e) Descent. Stability must be shown throughout the speed range 20 
knots either side of trim with--
    (1) The helicopter trimmed at 0.8 VH or 0.8 
VNEI (or 0.8 VLE for the landing gear extended 
case), whichever is lower;
    (2) Power required for 1,000 fpm descent at trim speed; and
    (3) Landing gear extended and retracted, if applicable.
    (f) Approach. Stability must be shown throughout the speed range 
from 0.7 times the minimum recommended approach speed to 20 knots above 
the maximum recommended approach speed with--
    (1) The helicopter trimmed at the recommended approach speed or 
speeds;
    (2) Landing gear extended and retracted, if applicable; and
    (3) Power required to maintain a 3[deg] glide path and power 
required to maintain the steepest approach gradient for which approval 
is requested.
    V. Static Lateral Directional Stability
    (a) Static directional stability must be positive throughout the 
approved ranges of airspeed, power, and vertical speed. In straight and 
steady sideslips up to 10[deg] from trim, 
directional control position must increase without discontinuity with 
the angle of sideslip, except for a small range of sideslip angles 
around trim. At greater angles up to the maximum sideslip angle 
appropriate to the type, increased directional control position must 
produce an increased angle of sideslip. It must be possible to maintain 
balanced flight without exceptional pilot skill or alertness.
    (b) During sideslips up to 10[deg] from trim 
throughout the approved ranges of airspeed, power, and vertical speed 
there must be no negative dihedral stability perceptible to the pilot 
through lateral control motion or force. Longitudinal cyclic movement 
with sideslip must not be excessive.
    VI. Dynamic stability. (a) Any oscillation having a period of less 
than 5 seconds must damp to \1/2\ amplitude in not more than one cycle.
    (b) Any oscillation having a period of 5 seconds or more but less 
than 10 seconds must damp to \1/2\ amplitude in not more than two 
cycles.
    (c) Any oscillation having a period of 10 seconds or more but less 
than 20 seconds must be damped.
    (d) Any oscillation having a period of 20 seconds or more may not 
achieve double amplitude in less than 20 seconds.

[[Page 852]]

    (e) Any aperiodic response may not achieve double amplitude in less 
than 9 seconds.
    VII. Stability Augmentation System (SAS)
    (a) If a SAS is used, the reliability of the SAS must be related to 
the effects of its failure. Any SAS failure condition that would prevent 
continued safe flight and landing must be extremely improbable. It must 
be shown that, for any failure condition of the SAS that is not shown to 
be extremely improbable--
    (1) The helicopter is safely controllable when the failure or 
malfunction occurs at any speed or altitude within the approved IFR 
operating limitations; and
    (2) The overall flight characteristics of the helicopter allow for 
prolonged instrument flight without undue pilot effort. Additional 
unrelated probable failures affecting the control system must be 
considered. In addition--
    (i) The controllability and maneuverability requirements in Subpart 
B must be met throughout a practical flight envelope;
    (ii) The flight control, trim, and dynamic stability characteristics 
must not be impaired below a level needed to allow continued safe flight 
and landing;
    (iii) For Category A helicopters, the dynamic stability requirements 
of Subpart B must also be met throughout a practical flight envelope; 
and
    (iv) The static longitudinal and static directional stability 
requirements of Subpart B must be met throughout a practical flight 
envelope.
    (b) The SAS must be designed so that it cannot create a hazardous 
deviation in flight path or produce hazardous loads on the helicopter 
during normal operation or in the event of malfunction or failure, 
assuming corrective action begins within an appropriate period of time. 
Where multiple systems are installed, subsequent malfunction conditions 
must be considered in sequence unless their occurrence is shown to be 
improbable.
    VIII. Equipment, systems, and installation. The basic equipment and 
installation must comply with Subpart F of Part 29 through Amendment 29-
14, with the following exceptions and additions:
    (a) Flight and navigation instruments. (1) A magnetic gyro-
stabilized direction indicator instead of the gyroscopic direction 
indicator required by Sec. 29.1303(h); and
    (2) A standby attitude indicator which meets the requirements of 
Sec. Sec. 29.1303(g)(1) through (7), instead of a rate-of-turn 
indicator required by Sec. 29.1303(g). If standby batteries are 
provided, they may be charged from the aircraft electrical system if 
adequate isolation is incorporated. The system must be designed so that 
the standby batteries may not be used for engine starting.
    (b) Miscellaneous requirements. (1) Instrument systems and other 
systems essential for IFR flight that could be adversely affected by 
icing must be provided with adequate ice protection whether or not the 
rotorcraft is certificated for operation in icing conditions.
    (2) There must be means in the generating system to automatically 
de-energize and disconnect from the main bus any power source developing 
hazardous overvoltage.
    (3) Each required flight instrument using a power supply (electric, 
vacuum, etc.) must have a visual means integral with the instrument to 
indicate the adequacy of the power being supplied.
    (4) When multiple systems performing like functions are required, 
each system must be grouped, routed, and spaced so that physical 
separation between systems is provided to ensure that a single 
malfunction will not adversely affect more than one system.
    (5) For systems that operate the required flight instruments at each 
pilot's station--
    (i) Only the required flight instruments for the first pilot may be 
connected to that operating system;
    (ii) Additional instruments, systems, or equipment may not be 
connected to an operating system for a second pilot unless provisions 
are made to ensure the continued normal functioning of the required 
instruments in the event of any malfunction of the additional 
instruments, systems, or equipment which is not shown to be extremely 
improbable;
    (iii) The equipment, systems, and installations must be designed so 
that one display of the information essential to the safety of flight 
which is provided by the instruments will remain available to a pilot, 
without additional crew-member action, after any single failure or 
combination of failures that is not shown to be extremely improbable; 
and
    (iv) For single-pilot configurations, instruments which require a 
static source must be provided with a means of selecting an alternate 
source and that source must be calibrated.
    (6) In determining compliance with the requirements of Sec. 
29.1351(d)(2), the supply of electrical power to all systems necessary 
for flight under IFR must be included in the evaluation.
    (c) Thunderstorm lights. In addition to the instrument lights 
required by Sec. 29.1381(a), thunderstorm lights which provide high 
intensity white flood lighting to the basic flight instruments must be 
provided. The thunderstorm lights must be installed to meet the 
requirements of Sec. 29.1381(b).
    IX. Rotorcraft Flight Manual. A Rotorcraft Flight Manual or 
Rotorcraft Flight Manual IFR Supplement must be provided and must 
contain--
    (a) Limitations. The approved IFR flight envelope, the IFR 
flightcrew composition, the revised kinds of operation, and the steepest

[[Page 853]]

IFR precision approach gradient for which the helicopter is approved;
    (b) Procedures. Required information for proper operation of IFR 
systems and the recommended procedures in the event of stability 
augmentation or electrical system failures; and
    (c) Performance. If VYI differs from VY, climb 
performance at VYI and with maximum continuous power 
throughout the ranges of weight, altitude, and temperature for which 
approval is requested.

[Amdt. 29-21, 48 FR 4392, Jan. 31, 1983, as amended by Amdt. 29-31, 55 
FR 38967, Sept. 21, 1990; 55 FR 41309, Oct. 10, 1990; Amdt. 29-40, 61 FR 
21908, May 10, 1996; Amdt. 29-51, 73 FR 11002, Feb. 29, 2008]



             Sec. Appendix C to Part 29--Icing Certification

    (a) Continuous maximum icing. The maximum continuous intensity of 
atmospheric icing conditions (continuous maximum icing) is defined by 
the variables of the cloud liquid water content, the mean effective 
diameter of the cloud droplets, the ambient air temperature, and the 
interrelationship of these three variables as shown in Figure 1 of this 
appendix. The limiting icing envelope in terms of altitude and 
temperature is given in Figure 2 of this appendix. The interrelationship 
of cloud liquid water content with drop diameter and altitude is 
determined from Figures 1 and 2. The cloud liquid water content for 
continuous maximum icing conditions of a horizontal extent, other than 
17.4 nautical miles, is determined by the value of liquid water content 
of Figure 1, multiplied by the appropriate factor from Figure 3 of this 
appendix.
    (b) Intermittent maximum icing. The intermittent maximum intensity 
of atmospheric icing conditions (intermittent maximum icing) is defined 
by the variables of the cloud liquid water content, the mean effective 
diameter of the cloud droplets, the ambient air temperature, and the 
interrelationship of these three variables as shown in Figure 4 of this 
appendix. The limiting icing envelope in terms of altitude and 
temperature is given in Figure 5 of this appendix. The interrelationship 
of cloud liquid water content with drop diameter and altitude is 
determined from Figures 4 and 5. The cloud liquid water content for 
intermittent maximum icing conditions of a horizontal extent, other than 
2.6 nautical miles, is determined by the value of cloud liquid water 
content of Figure 4 multiplied by the appropriate factor in Figure 6 of 
this appendix.

[[Page 854]]

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[[Page 855]]


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[[Page 856]]


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[[Page 857]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.094


[[Page 858]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.095


[[Page 859]]


[GRAPHIC] [TIFF OMITTED] TC28SE91.096


[Amdt. 29-21, 48 FR 4393, Jan. 31, 1983]

[[Page 860]]



  Sec. Appendix D to Part 29--Criteria for Demonstration of Emergency 
                Evacuation Procedures Under Sec. 29.803

    (a) The demonstration must be conducted either during the dark of 
the night or during daylight with the dark of night simulated. If the 
demonstration is conducted indoors during daylight hours, it must be 
conducted inside a darkened hangar having doors and windows covered. In 
addition, the doors and windows of the rotorcraft must be covered if the 
hangar illumination exceeds that of a moonless night. Illumination on 
the floor or ground may be used, but it must be kept low and shielded 
against shining into the rotorcraft's windows or doors.
    (b) The rotorcraft must be in a normal attitude with landing gear 
extended.
    (c) Safety equipment such as mats or inverted liferafts may be 
placed on the floor or ground to protect participants. No other 
equipment that is not part of the rotorcraft's emergency evacuation 
equipment may be used to aid the participants in reaching the ground.
    (d) Except as provided in paragraph (a) of this appendix, only the 
rotorcraft's emergency lighting system may provide illumination.
    (e) All emergency equipment required for the planned operation of 
the rotorcraft must be installed.
    (f) Each external door and exit and each internal door or curtain 
must be in the takeoff configuration.
    (g) Each crewmember must be seated in the normally assigned seat for 
takeoff and must remain in that seat until receiving the signal for 
commencement of the demonstration. For compliance with this section, 
each crewmember must be--
    (1) A member of a regularly scheduled line crew; or
    (2) A person having knowledge of the operation of exits and 
emergency equipment.
    (h) A representative passenger load of persons in normal health must 
be used as follows:
    (1) At least 25 percent must be over 50 years of age, with at least 
40 percent of these being females.
    (2) The remaining, 75 percent or less, must be 50 years of age or 
younger, with at least 30 percent of these being females.
    (3) Three life-size dolls, not included as part of the total 
passenger load, must be carried by passengers to simulate live infants 2 
years old or younger, except for a total passenger load of fewer than 44 
but more than 19, one doll must be carried. A doll is not required for a 
19 or fewer passenger load.
    (4) Crewmembers, mechanics, and training personnel who maintain or 
operate the rotorcraft in the normal course of their duties may not be 
used as passengers.
    (i) No passenger may be assigned a specific seat except as the 
Administrator may require. Except as required by paragraph (1) of this 
appendix, no employee of the applicant may be seated next to an 
emergency exit, except as allowed by the Administrator.
    (j) Seat belts and shoulder harnesses (as required) must be 
fastened.
    (k) Before the start of the demonstration, approximately one-half of 
the total average amount of carry-on baggage, blankets, pillows, and 
other similar articles must be distributed at several locations in the 
aisles and emergency exit access ways to create minor obstructions.
    (l) No prior indication may be given to any crewmember or passenger 
of the particular exits to be used in the demonstration.
    (m) The applicant may not practice, rehearse, or describe the 
demonstration for the participants nor may any participant have taken 
part in this type of demonstration within the preceding 6 months.
    (n) A pretakeoff passenger briefing may be given. The passengers may 
also be advised to follow directions of crewmembers, but not be 
instructed on the procedures to be followed in the demonstration.
    (o) If safety equipment, as allowed by paragraph (c) of this 
appendix, is provided, either all passenger and cockpit windows must be 
blacked out or all emergency exits must have safety equipment to prevent 
disclosure of the available emergency exits.
    (p) Not more than 50 percent of the emergency exits in the sides of 
the fuselage of a rotorcraft that meet all of the requirements 
applicable to the required emergency exits for that rotorcraft may be 
used for demonstration. Exits that are not to be used for the 
demonstration must have the exit handle deactivated or must be indicated 
by red lights, red tape, or other acceptable means placed outside the 
exits to indicate fire or other reasons why they are unusable. The exits 
to be used must be representative of all the emergency exits on the 
rotorcraft and must be designated by the applicant, subject to approval 
by the Administrator. If installed, at least one floor level exit (Type 
I; Sec. 29.807(a)(1)) must be used as required by Sec. 29.807(c).
    (q) All evacuees must leave the rotorcraft by a means provided as 
part of the rotorcraft's equipment.
    (r) Approved procedures must be fully utilized during the 
demonstration.
    (s) The evacuation time period is completed when the last occupant 
has evacuated the rotorcraft and is on the ground.

[Amdt. 27-26, 55 FR 8005, Mar. 6, 1990]

[[Page 861]]



 Sec. Appendix E to Part 29--HIRF Environments and Equipment HIRF Test 
                                 Levels

    This appendix specifies the HIRF environments and equipment HIRF 
test levels for electrical and electronic systems under Sec. 29.1317. 
The field strength values for the HIRF environments and laboratory 
equipment HIRF test levels are expressed in root-mean-square units 
measured during the peak of the modulation cycle.
    (a) HIRF environment I is specified in the following table:

                      Table I.--HIRF Environment I
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-2 MHz......................................         50         50
2 MHz-30 MHz......................................        100        100
30 MHz-100 MHz....................................         50         50
100 MHz-400 MHz...................................        100        100
400 MHz-700 MHz...................................        700         50
700 MHz-1 GHz.....................................        700        100
1 GHz-2 GHz.......................................      2,000        200
2 GHz-6 GHz.......................................      3,000        200
6 GHz-8 GHz.......................................      1,000        200
8 GHz-12 GHz......................................      3,000        300
12 GHz-18 GHz.....................................      2,000        200
18 GHz-40 GHz.....................................        600        200
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (b) HIRF environment II is specified in the following table:

                     Table II.--HIRF Environment II
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-500 kHz....................................         20         20
500 kHz-2 MHz.....................................         30         30
2 MHz-30 MHz......................................        100        100
30 MHz-100 MHz....................................         10         10
100 MHz-200 MHz...................................         30         10
200 MHz-400 MHz...................................         10         10
400 MHz-1 GHz.....................................        700         40
1 GHz-2 GHz.......................................      1,300        160
2 GHz-4 GHz.......................................      3,000        120
4 GHz-6 GHz.......................................      3,000        160
6 GHz-8 GHz.......................................        400        170
8 GHz-12 GHz......................................      1,230        230
12 GHz-18 GHz.....................................        730        190
18 GHz-40 GHz.....................................        600        150
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (c) HIRF environment III is specified in the following table:

                    Table III.--HIRF Environment III
------------------------------------------------------------------------
                                                       Field strength
                                                        (volts/meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-100 kHz....................................        150        150
100 kHz-400 MHz...................................        200        200
400 MHz-700 MHz...................................        730        200
700 MHz-1 GHz.....................................      1,400        240
1 GHz-2 GHz.......................................      5,000        250
2 GHz-4 GHz.......................................      6,000        490
4 GHz-6 GHz.......................................      7,200        400
6 GHz-8 GHz.......................................      1,100        170
8 GHz-12 GHz......................................      5,000        330
12 GHz-18 GHz.....................................      2,000        330
18 GHz-40 GHz.....................................      1,000        420
------------------------------------------------------------------------
In this table, the higher field strength applies at the frequency band
  edges.

    (d) Equipment HIRF Test Level 1. (1) From 10 kilohertz (kHz) to 400 
megahertz (MHz), use conducted susceptibility tests with continuous wave 
(CW) and 1 kHz square wave modulation with 90 percent depth or greater. 
The conducted susceptibility current must start at a minimum of 0.6 
milliamperes (mA) at 10 kHz, increasing 20 decibel (dB) per frequency 
decade to a minimum of 30 mA at 500 kHz.
    (2) From 500 kHz to 40 MHz, the conducted susceptibility current 
must be at least 30 mA.
    (3) From 40 MHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 30 mA at 40 MHz, decreasing 20 dB per frequency 
decade to a minimum of 3 mA at 400 MHz.
    (4) From 100 MHz to 400 MHz, use radiated susceptibility tests at a 
minimum of 20 volts per meter (V/m) peak with CW and 1 kHz square wave 
modulation with 90 percent depth or greater.
    (5) From 400 MHz to 8 gigahertz (GHz), use radiated susceptibility 
tests at a minimum of 150 V/m peak with pulse modulation of 4 percent 
duty cycle with a 1 kHz pulse repetition frequency. This signal must be 
switched on and off at a rate of 1 Hz with a duty cycle of 50 percent.
    (e) Equipment HIRF Test Level 2. Equipment HIRF test level 2 is HIRF 
environment II in table II of this appendix reduced by acceptable 
aircraft transfer function and attenuation curves. Testing must cover 
the frequency band of 10 kHz to 8 GHz.
    (f) Equipment HIRF Test Level 3. (1) From 10 kHz to 400 MHz, use 
conducted susceptibility tests, starting at a minimum of 0.15 mA at 10 
kHz, increasing 20 dB per frequency decade to a minimum of 7.5 mA at 500 
kHz.
    (2) From 500 kHz to 40 MHz, use conducted susceptibility tests at a 
minimum of 7.5 mA.
    (3) From 40 MHz to 400 MHz, use conducted susceptibility tests, 
starting at a minimum of 7.5 mA at 40 MHz, decreasing 20 dB per 
frequency decade to a minimum of 0.75 mA at 400 MHz.

[[Page 862]]

    (4) From 100 MHz to 8 GHz, use radiated susceptibility tests at a 
minimum of 5 V/m.

[Doc. No. FAA-2006-23657, 72 FR 44028, Aug. 6, 2007]



PART 31_AIRWORTHINESS STANDARDS: MANNED FREE BALLOONS--Table of Contents



                            Subpart A_General

Sec.
31.1 Applicability.

                      Subpart B_Flight Requirements

31.12 Proof of compliance.
31.14 Weight limits.
31.16 Empty weight.
31.17 Performance: Climb.
31.19 Performance: Uncontrolled descent.
31.20 Controllability.

                     Subpart C_Strength Requirements

31.21 Loads.
31.23 Flight load factor.
31.25 Factor of safety.
31.27 Strength.

                      Subpart D_Design Construction

31.31 General.
31.33 Materials.
31.35 Fabrication methods.
31.37 Fastenings.
31.39 Protection.
31.41 Inspection provisions.
31.43 Fitting factor.
31.45 Fuel cells.
31.46 Pressurized fuel systems.
31.47 Burners.
31.49 Control systems.
31.51 Ballast.
31.53 Drag rope.
31.55 Deflation means.
31.57 Rip cords.
31.59 Trapeze, basket, or other means provided for occupants.
31.61 Static discharge.
31.63 Safety belts.
31.65 Position lights.

                           Subpart E_Equipment

31.71 Function and installation.

             Subpart F_Operating Limitations and Information

31.81 General.
31.82 Instructions for Continued Airworthiness.
31.83 Conspicuity.
31.85 Required basic equipment.

Appendix A to Part 31--Instructions for Continued Airworthiness

    Authority: 49 U.S.C. 106(g), 40113, 44701-44702, 44704.

    Source: Docket No. 1437, 29 FR 8258, July 1, 1964, as amended by 
Amdt. 31-1, 29 FR 14563, Oct. 24, 1964, unless otherwise noted.



                            Subpart A_General



Sec. 31.1  Applicability.

    (a) This part prescribes airworthiness standards for the issue of 
type certificates and changes to those certificates, for manned free 
balloons.
    (b) Each person who applies under Part 21 for such a certificate or 
change must show compliance with the applicable requirements of this 
part.
    (c) For purposes of this part--
    (1) A captive gas balloon is a balloon that derives its lift from a 
captive lighter-than-air gas;
    (2) A hot air balloon is a balloon that derives its lift from heated 
air;
    (3) The envelope is the enclosure in which the lifting means is 
contained;
    (4) The basket is the container, suspended beneath the envelope, for 
the balloon occupants;
    (5) The trapeze is a harness or is a seat consisting of a horizontal 
bar or platform suspended beneath the envelope for the balloon 
occupants; and
    (6) The design maximum weight is the maximum total weight of the 
balloon, less the lifting gas or air.

[Doc. No. 1437, 29 FR 8258, July 1, 1964, as amended by Amdt. 31-3, 41 
FR 55474, Dec. 20, 1976]



                      Subpart B_Flight Requirements



Sec. 31.12  Proof of compliance.

    (a) Each requirement of this subpart must be met at each weight 
within the range of loading conditions for which certification is 
requested. This must be shown by--
    (1) Tests upon a balloon of the type for which certification is 
requested or by calculations based on, and equal in accuracy to, the 
results of testing; and
    (2) Systematic investigation of each weight if compliance cannot be 
reasonably inferred from the weights investigated.

[[Page 863]]

    (b) Except as provided in Sec. 31.17(b), allowable weight 
tolerances during flight testing are +5 percent and -10 percent.

[Amdt. 31-4, 45 FR 60179, Sept. 11, 1980]



Sec. 31.14  Weight limits.

    (a) The range of weights over which the balloon may be safely 
operated must be established.
    (b) Maximum weight. The maximum weight is the highest weight at 
which compliance with each applicable requirement of this part is shown. 
The maximum weight must be established so that it is not more than--
    (1) The highest weight selected by the applicant;
    (2) The design maximum weight which is the highest weight at which 
compliance with each applicable structural loading condition of this 
part is shown; or
    (3) The highest weight at which compliance with each applicable 
flight requirement of this part is shown.
    (c) The information established under paragraphs (a) and (b) of this 
section must be made available to the pilot in accordance with Sec. 
31.81.

[Amdt. 31-3, 41 FR 55474, Dec. 20, 1976]



Sec. 31.16  Empty weight.

    The empty weight must be determined by weighing the balloon with 
installed equipment but without lifting gas or heater fuel.

[Amdt. 31-4, 45 FR 60179, Sept. 11, 1980]



Sec. 31.17  Performance: Climb.

    (a) Each balloon must be capable of climbing at least 300 feet in 
the first minute after takeoff with a steady rate of climb. Compliance 
with the requirements of this section must be shown at each altitude and 
ambient temperature for which approval is sought.
    (b) Compliance with the requirements of paragraph (a) of this 
section must be shown at the maximum weight with a weight tolerance of 
+5 percent.

[Amdt. 31-4, 45 FR 60179, Sept. 11, 1980]



Sec. 31.19  Performance: Uncontrolled descent.

    (a) The following must be determined for the most critical 
uncontrolled descent that can result from any single failure of the 
heater assembly, fuel cell system, gas value system, or maneuvering vent 
system, or from any single tear in the balloon envelope between tear 
stoppers:
    (1) The maximum vertical velocity attained.
    (2) The altitude loss from the point of failure to the point at 
which maximum vertical velocity is attained.
    (3) The altitude required to achieve level flight after corrective 
action is inititated, with the balloon descending at the maximum 
vertical velocity determined in paragraph (a)(1) of this section.
    (b) Procedures must be established for landing at the maximum 
vertical velocity determined in paragraph (a)(1) of this section and for 
arresting that descent rate in accordance with paragraph (a)(3) of this 
section.

[Amdt. 31-4, 45 FR 60179, Sept. 11, 1980]



Sec. 31.20  Controllability.

    The applicant must show that the balloon is safely controllable and 
maneuverable during takeoff, ascent, descent, and landing without 
requiring exceptional piloting skill.

[Amdt. 31-3, 41 FR 55474, Dec. 20, 1976]



                     Subpart C_Strength Requirements



Sec. 31.21  Loads.

    Strength requirements are specified in terms of limit loads, that 
are the maximum load to be expected in service, and ultimate loads, that 
are limit loads multiplied by prescribed factors of safety. Unless 
otherwise specified, all prescribed loads are limit loads.



Sec. 31.23  Flight load factor.

    In determining limit load, the limit flight load factor must be at 
least 1.4.



Sec. 31.25  Factor of safety.

    (a) Except as specified in paragraphs (b) and (c) of this section, 
the factor of safety is 1.5.
    (b) A factor of safety of at least five must be used in envelope 
design. A reduced factor of safety of at least two may be used if it is 
shown that the selected factor will preclude failure due to creep or 
instantaneous rupture from

[[Page 864]]

lack of rip stoppers. The selected factor must be applied to the more 
critical of the maximum operating pressure or envelope stress.
    (c) A factor of safety of at least five must be used in the design 
of all fibrous or non-metallic parts of the rigging and related 
attachments of the envelope to basket, trapeze, or other means provided 
for carrying occupants. The primary attachments of the envelope to the 
basket, trapeze, or other means provided for carrying occupants must be 
designed so that failure is extremely remote or so that any single 
failure will not jeopardize safety of flight.
    (d) In applying factors of safety, the effect of temperature, and 
other operating characteristics, or both, that may affect strength of 
the balloon must be accounted for.
    (e) For design purposes, an occupant weight of at least 170 pounds 
must be assumed.

[Doc. No. 1437, 29 FR 8258, July 1, 1964, as amended by Amdt. 31-2, 30 
FR 3377, Mar. 13, 1965]



Sec. 31.27  Strength.

    (a) The structure must be able to support limit loads without 
detrimental effect.
    (b) The structure must be substantiated by test to be able to 
withstand the ultimate loads for at least three seconds without failure. 
For the envelope, a test of a representative part is acceptable, if the 
part tested is large enough to include critical seams, joints, and load 
attachment points and members.
    (c) An ultimate free-fall drop test must be made of the basket, 
trapeze, or other place provided for occupants. The test must be made at 
design maximum weight on a horizontal surface, with the basket, trapeze, 
or other means provided for carrying occupants, striking the surface at 
angles of 0, 15, and 30 degrees. The weight may be distributed to 
simulate actual conditions. There must be no distortion or failure that 
is likely to cause serious injury to the occupants. A drop test height 
of 36 inches, or a drop test height that produces, upon impact, a 
velocity equal to the maximum vertical velocity determined in accordance 
with Sec. 31.19, whichever is higher, must be used.

[Doc. No. 1437, 29 FR 8258, July 1, 1964, as amended by Amdt. 31-4, 45 
FR 60179, Sept. 11, 1980]



                      Subpart D_Design Construction



Sec. 31.31  General.

    The suitability of each design detail or part that bears on safety 
must be established by tests or analysis.



Sec. 31.33  Materials.

    (a) The suitability and durability of all materials must be 
established on the basis of experience or tests. Materials must conform 
to approved specifications that will ensure that they have the strength 
and other properties assumed in the design data.
    (b) Material strength properties must be based on enough tests of 
material conforming to specifications so as to establish design values 
on a statistical basis.



Sec. 31.35  Fabrication methods.

    The methods of fabrication used must produce a consistently sound 
structure. If a fabrication process requires close control to reach this 
objective, the process must be performed in accordance with an approved 
process specification.



Sec. 31.37  Fastenings.

    Only approved bolts, pins, screws, and rivets may be used in the 
structure. Approved locking devices or methods must be used for all 
these bolts, pins, and screws, unless the installation is shown to be 
free from vibration. Self-locking nuts may not be used on bolts that are 
subject to rotation in service.



Sec. 31.39  Protection.

    Each part of the balloon must be suitably protected against 
deterioration or loss of strength in service due to weathering, 
corrosion, or other causes.



Sec. 31.41  Inspection provisions.

    There must be a means to allow close examination of each part that 
require repeated inspection and adjustment.

[[Page 865]]



Sec. 31.43  Fitting factor.

    (a) A fitting factor of at least 1.15 must be used in the analysis 
of each fitting the strength of which is not proven by limit and 
ultimate load tests in which the actual stress conditions are simulated 
in the fitting and surrounding structure. This factor applies to all 
parts of the fitting, the means of attachment, and the bearing on the 
members joined.
    (b) Each part with an integral fitting must be treated as a fitting 
up to the point where the section properties become typical of the 
member.
    (c) The fitting factor need not be used if the joint design is made 
in accordance with approved practices and is based on comprehensive test 
data.



Sec. 31.45  Fuel cells.

    If fuel cells are used, the fuel cells, their attachments, and 
related supporting structure must be shown by tests to be capable of 
withstanding, without detrimental distortion or failure, any inertia 
loads to which the installation may be subjected, including the drop 
tests prescribed in Sec. 31.27(c). In the tests, the fuel cells must be 
loaded to the weight and pressure equivalent to the full fuel quantity 
condition.

[Amdt. 31-3, 41 FR 55474, Dec. 20, 1976]



Sec. 31.46  Pressurized fuel systems.

    For pressurized fuel systems, each element and its connecting 
fittings and lines must be tested to an ultimate pressure of at least 
twice the maximum pressure to which the system will be subjected in 
normal operation. No part of the system may fail or malfunction during 
the test. The test configuration must be representative of the normal 
fuel system installation and balloon configuration.

[Amdt. 31-3, 41 FR 55474, Dec. 20, 1976]



Sec. 31.47  Burners.

    (a) If a burner is used to provide the lifting means, the system 
must be designed and installed so as not to create a fire hazard.
    (b) There must be shielding to protect parts adjacent to the burner 
flame, and the occupants, from heat effects.
    (c) There must be controls, instruments, or other equipment 
essential to the safe control and operation of the heater. They must be 
shown to be able to perform their intended functions during normal and 
emergency operation.
    (d) The burner system (including the burner unit, controls, fuel 
lines, fuel cells, regulators, control valves, and other related 
elements) must be substantiated by an endurance test of at least 40 
hours. Each element of the system must be installed and tested to 
simulate actual balloon installation and use.
    (1) The test program for the main blast valve operation of the 
burner must include:
    (i) Five hours at the maximum fuel pressure for which approval is 
sought, with a burn time for each one minute cycle of three to ten 
seconds. The burn time must be established so that each burner is 
subjected to the maximum thermal shock for temperature affected 
elements;
    (ii) Seven and one-half hours at an intermediate fuel pressure, with 
a burn time for each one minute cycle of three to ten seconds. An 
intermediate fuel pressure is 40 to 60 percent of the range between the 
maximum fuel pressure referenced in paragraph (d)(1)(i) of this section 
and minimum fuel pressure referenced in paragraph (d)(1)(iii);
    (iii) Six hours and fifteen minutes at the minimum fuel pressure for 
which approval is sought, with a burn time for each one minute cycle of 
three to ten seconds;
    (iv) Fifteen minutes of operation on vapor, with a burn time for 
each one minute cycle of at least 30 seconds; and
    (v) Fifteen hours of normal flight operation.
    (2) The test program for the secondary or backup operation of the 
burner must include six hours of operation with a burn time for each 
five minute cycle of one minute at an intermediate fuel pressure.
    (e) The test must also include at least three flameouts and 
restarts.

[[Page 866]]

    (f) Each element of the system must be serviceable at the end of the 
test.

[Doc. No. 1437, 29 FR 8258, July 1, 1964, as amended by Amdt. 31-2, 30 
FR 3377, Mar. 13, 1965; Amdt. 31-7, 61 FR 18223, Apr. 24, 1996; 61 FR 
20877, May 8, 1996]



Sec. 31.49  Control systems.

    (a) Each control must operate easily, smoothly, and positively 
enough to allow proper performance of its functions. Controls must be 
arranged and identified to provide for convenience of operation and to 
prevent the possibility of confusion and subsequent inadvertent 
operation.
    (b) Each control system and operating device must be designed and 
installed in a manner that will prevent jamming, chafing, or 
interference from passengers, cargo, or loose objects. Precaution must 
be taken to prevent foreign objects from jamming the controls. The 
elements of the control system must have design features or must be 
distinctly and permanently marked to minimize the possibility of 
incorrect assembly that could result in malfunctioning of the control 
system.
    (c) Each balloon using a captive gas as the lifting means must have 
an automatic valve or appendix that is able to release gas automatically 
at the rate of at least three percent of the total volume per minute 
when the balloon is at its maximum operating pressure.
    (d) Each hot air balloon must have a means to allow the controlled 
release of hot air during flight.
    (e) Each hot air balloon must have a means to indicate the maximum 
envelope skin temperatures occurring during operation. The indicator 
must be readily visible to the pilot and marked to indicate the limiting 
safe temperature of the envelope material. If the markings are on the 
cover glass of the instrument, there must be provisions to maintain the 
correct alignment of the glass cover with the face of the dial.

[Doc. No. 1437, 29 FR 8258, July 1, 1964, as amended by Amdt. 31-2, 30 
FR 3377, Mar. 13, 1965]



Sec. 31.51  Ballast.

    Each captive gas balloon must have a means for the safe storage and 
controlled release of ballast. The ballast must consist of material 
that, if released during flight, is not hazardous to persons on the 
ground.



Sec. 31.53  Drag rope.

    If a drag rope is used, the end that is released overboard must be 
stiffened to preclude the probability of the rope becoming entangled 
with trees, wires, or other objects on the ground.



Sec. 31.55  Deflation means.

    There must be a means to allow emergency deflation of the envelope 
so as to allow a safe emergency landing. If a system other than a manual 
system is used, the reliability of the system used must be 
substantiated.

[Amdt. 31-2, 30 FR 3377, Mar. 13, 1965]



Sec. 31.57  Rip cords.

    (a) If a rip cord is used for emergency deflation, it must be 
designed and installed to preclude entanglement.
    (b) The force required to operate the rip cord may not be less than 
25, or more than 75, pounds.
    (c) The end of the rip cord to be operated by the pilot must be 
colored red.
    (d) The rip cord must be long enough to allow an increase of at 
least 10 percent in the vertical dimension of the envelope.



Sec. 31.59  Trapeze, basket, or other means provided for occupants.

    (a) The trapeze, basket, or other means provided for carrying 
occupants may not rotate independently of the envelope.
    (b) Each projecting object on the trapeze, basket, or other means 
provided for carrying occupants, that could cause injury to the 
occupants, must be padded.



Sec. 31.61  Static discharge.

    Unless shown not to be necessary for safety, there must be 
appropriate bonding means in the design of each balloon using flammable 
gas as a lifting means to ensure that the effects of static discharges 
will not create a hazard.

[Amdt. 31-2, 30 FR 3377, Mar. 13, 1965]

[[Page 867]]



Sec. 31.63  Safety belts.

    (a) There must be a safety belt, harness, or other restraining means 
for each occupant, unless the Administrator finds it unnecessary. If 
installed, the belt, harness, or other restraining means and its 
supporting structure must meet the strength requirements of subpart C of 
this part.
    (b) This section does not apply to balloons that incorporate a 
basket or gondola.

[Amdt. 31-2, 30 FR 3377, Mar. 13, 1965, as amended by Amdt. 31-3, 41 FR 
55474, Dec. 20, 1976]



Sec. 31.65  Position lights.

    (a) If position lights are installed, there must be one steady 
aviation white position light and one flashing aviation red (or flashing 
aviation white) position light with an effective flash frequency of at 
least 40, but not more than 100, cycles per minute.
    (b) Each light must provide 360[deg] horizontal coverage at the 
intensities prescribed in this paragraph. The following light 
intensities must be determined with the light source operating at a 
steady state and with all light covers and color filters in place and at 
the manufacturer's rated mimimum voltage. For the flashing aviation red 
light, the measured values must be adjusted to correspond to a red 
filter temperature of at least 130 [deg]F:
    (1) The intensities in the horizontal plane passing through the 
light unit must equal or exceed the following values:

------------------------------------------------------------------------
                                                                Minimum
                       Position light                          intensity
                                                               (candles)
------------------------------------------------------------------------
Steady white................................................          20
Flashing red or white.......................................          40
------------------------------------------------------------------------

    (2) The intensities in vertical planes must equal or exceed the 
following values. An intensity of one unit corresponds to the applicable 
horizontal plane intensity specified in paragraph (b)(1) of this 
section.

------------------------------------------------------------------------
                                                                Minimum
 Angles above and below the horizontal in any vertical plane   intensity
                          (degrees)                             (units)
------------------------------------------------------------------------
0...........................................................        1.00
0 to 5......................................................        0.90
5 to 10.....................................................        0.80
10 to 15....................................................        0.70
15 to 20....................................................        0.50
20 to 30....................................................        0.30
30 to 40....................................................        0.10
40 to 60....................................................        0.05
------------------------------------------------------------------------

    (c) The steady white light must be located not more than 20 feet 
below the basket, trapeze, or other means for carrying occupants. The 
flashing red or white light must be located not less than 7, nor more 
than 10, feet below the steady white light.
    (d) There must be a means to retract and store the lights.
    (e) Each position light color must have the applicable International 
Commission on Illumination chromaticity coordinates as follows:
    (1) Aviation red--

y is not greater than 0.335; and z is not greater than 0.002.

    (2) Aviation white--

x is not less than 0.300 and not greater than 0.540;
y is not less than x-0.040 or yo-0.010, whichever is the 
          smaller; and
y is not greater than x+0.020 nor 0.636-0.0400 x;
Where yo is the y coordinate of the Planckian radiator for 
          the value of x considered.

[Doc. No. 1437, 29 FR 8258, July 1, 1964, as amended by Amdt. 31-1, 29 
FR 14563, Oct. 24, 1964; Amdt. 31-4, 45 FR 60179, Sept. 11, 1980]



                           Subpart E_Equipment



Sec. 31.71  Function and installation.

    (a) Each item of installed equipment must--
    (1) Be of a kind and design appropriate to its intended function;
    (2) Be permanently and legibly marked or, if the item is too small 
to mark, tagged as to its identification, function, or operating 
limitations, or any applicable combination of those factors;
    (3) Be installed according to limitations specified for that 
equipment; and
    (4) Function properly when installed.
    (b) No item of installed equipment, when performing its function, 
may affect the function of any other equipment so as to create an unsafe 
condition.

[[Page 868]]

    (c) The equipment, systems, and installations must be designed to 
prevent hazards to the balloon in the event of a probable malfunction or 
failure.

[Amdt. 31-4, 45 FR 60180, Sept. 11, 1980]



             Subpart F_Operating Limitations and Information



Sec. 31.81  General.

    (a) The following information must be established:
    (1) Each operating limitation, including the maximum weight 
determined under Sec. 31.14.
    (2) The normal and emergency procedures.
    (3) Other information necessary for safe operation, including--
    (i) The empty weight determined under Sec. 31.16;
    (ii) The rate of climb determined under Sec. 31.17, and the 
procedures and conditions used to determine performance;
    (iii) The maximum vertical velocity, the altitude drop required to 
attain that velocity, and altitude drop required to recover from a 
descent at that velocity, determined under Sec. 31.19, and the 
procedures and conditions used to determine performance; and
    (iv) Pertinent information peculiar to the balloon's operating 
characteristics.
    (b) The information established in compliance with paragraph (a) of 
this section must be furnished by means of--
    (1) A Balloon Flight Manual; or
    (2) A placard on the balloon that is clearly visible to the pilot.

[Amdt. 31-4, 45 FR 60180, Sept. 11, 1980]



Sec. 31.82  Instructions for Continued Air worthiness.

    The applicant must prepare Instructions for Continued Airworthiness 
in accordance with appendix A to this part that are acceptable to the 
Administrator. The instructions may be incomplete at type certification 
if a program exists to ensure their completion prior to delivery of the 
first balloon or issuance of a standard certificate of airworthiness, 
whichever occurs later.

[Amdt. 31-4, 45 FR 60180, Sept. 11, 1980]



Sec. 31.83  Conspicuity.

    The exterior surface of the envelope must be of a contrasting color 
or colors so that it will be conspicuous during operation. However, 
multicolored banners or streamers are acceptable if it can be shown that 
they are large enough, and there are enough of them of contrasting 
color, to make the balloon conspicuous during flight.



Sec. 31.85  Required basic equipment.

    In addition to any equipment required by this subchapter for a 
specific kind of operation, the following equipment is required:
    (a) For all balloons:
    (1) [Reserved]
    (2) An altimeter.
    (3) A rate of climb indicator.
    (b) For hot air balloons:
    (1) A fuel quantity gauge. If fuel cells are used, means must be 
incorporated to indicate to the crew the quantity of fuel in each cell 
during flight. The means must be calibrated in appropriate units or in 
percent of fuel cell capacity.
    (2) An envelope temperature indicator.
    (c) For captive gas balloons, a compass.

[Amdt. 31-2, 30 FR 3377, Mar. 13, 1965, as amended by Amdt. 31-3, 41 FR 
55474, Dec. 20, 1976; Amdt. 31-4, 45 FR 60180, Sept. 11, 1980]



  Sec. Appendix A to Part 31--Instructions for Continued Airworthiness

                              a31.1 general

    (a) This appendix specifies requirements for the preparation of 
Instructions for Continued Airworthiness as required by Sec. 31.82.
    (b) The Instructions for Continued Airworthiness for each balloon 
must include the Instructions for Continued Airworthiness for all 
balloon parts required by this chapter and any required information 
relating to the interface of those parts with the balloon. If 
Instructions for Continued Airworthiness are not supplied by the part 
manufacturer for a balloon part, the Instructions for Continued 
Airworthiness for the balloon must include the information essential to 
the continued airworthiness of the balloon.
    (c) The applicant must submit to the FAA a program to show how 
changes to the Instructions for Continued Airworthiness made by the 
applicant or by the manufacturers of balloon parts will be distributed.

[[Page 869]]

                              a31.2 format

    (a) The Instructions for Continued Airworthiness must be in the form 
of a manual or manuals as appropriate for the quantity of data to be 
provided.
    (b) The format of the manual or manuals must provide for a practical 
arrangement.

                              a31.3 content

    The contents of the manual or manuals must be prepared in the 
English language. The Instructions for Continued Airworthiness must 
contain the following information:
    (a) Introduction information that includes an explanation of the 
balloon's features and data to the extent necessary for maintenance or 
preventive maintenance.
    (b) A description of the balloon and its systems and installations.
    (c) Basic control and operation information for the balloon and its 
components and systems.
    (d) Servicing information that covers details regarding servicing of 
balloon components, including burner nozzles, fuel tanks, and valves 
during operations.
    (e) Maintenance information for each part of the balloon and its 
envelope, controls, rigging, basket structure, fuel systems, 
instruments, and heater assembly that provides the recommended periods 
at which they should be cleaned, adjusted, tested, and lubricated, the 
applicable wear tolerances, and the degree of work recommended at these 
periods. However, the applicant may refer to an accessory, instrument, 
or equipment manufacturer as the source of this information if the 
applicant shows that the item has an exceptionally high degree of 
complexity requiring specialized maintenance techniques, test equipment, 
or expertise. The recommended overhaul periods and necessary cross 
references to the Airworthiness Limitations section of the manual must 
also be included. In addition, the applicant must include an inspection 
program that includes the frequency and extent of the inspections 
necessary to provide for the continued airworthiness of the balloon.
    (f) Troubleshooting information describing probable malfunctions, 
how to recognize those malfunctions, and the remedial action for those 
malfunctions.
    (g) Details of what, and how, to inspect after a hard landing.
    (h) Instructions for storage preparation including any storage 
limits.
    (i) Instructions for repair on the balloon envelope and its basket 
or trapeze.

                 a31.4 airworthiness limitations section

    The Instructions for Continued Airworthiness must contain a section 
titled Airworthiness Limitations that is segregated and clearly 
distinguishable from the rest of the document. This section must set 
forth each mandatory replacement time, structural inspection interval, 
and related structural inspection procedure, including envelope 
structural integrity, required for type certification. If the 
Instructions for Continued Airworthiness consist of multiple documents, 
the section required by this paragraph must be included in the principal 
manual. This section must contain a legible statement in a prominent 
location that reads: ``The Airworthiness Limitations section is FAA 
approved and specifies maintenance required under Sec. Sec. 43.16 and 
91.403 of the Federal Aviation Regulations.''

[Amdt. 31-4, 45 FR 60180, Sept. 11, 1980, as amended by Amdt. 31-5, 54 
FR 34330, Aug. 18, 1989]



PART 33_AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES--Table of Contents



                            Subpart A_General

Sec.
33.1 Applicability.
33.3 General.
33.4 Instructions for Continued Airworthiness.
33.5 Instruction manual for installing and operating the engine.
33.7 Engine ratings and operating limitations.
33.8 Selection of engine power and thrust ratings.

               Subpart B_Design and Construction; General

33.11 Applicability.
33.13 [Reserved]
33.15 Materials.
33.17 Fire protection.
33.19 Durability.
33.21 Engine cooling.
33.23 Engine mounting attachments and structure.
33.25 Accessory attachments.
33.27 Turbine, compressor, fan, and turbosupercharger rotor overspeed.
33.28 Engine control systems.
33.29 Instrument connection.

    Subpart C_Design and Construction; Reciprocating Aircraft Engines

33.31 Applicability.
33.33 Vibration.
33.34 Turbocharger rotors.
33.35 Fuel and induction system.
33.37 Ignition system.
33.39 Lubrication system.

          Subpart D_Block Tests; Reciprocating Aircraft Engines

33.41 Applicability.

[[Page 870]]

33.42 General.
33.43 Vibration test.
33.45 Calibration tests.
33.47 Detonation test.
33.49 Endurance test.
33.51 Operation test.
33.53 Engine system and component tests.
33.55 Teardown inspection.
33.57 General conduct of block tests.

       Subpart E_Design and Construction; Turbine Aircraft Engines

33.61 Applicability.
33.62 Stress analysis.
33.63 Vibration.
33.64 Pressurized engine static parts.
33.65 Surge and stall characteristics.
33.66 Bleed air system.
33.67 Fuel system.
33.68 Induction system icing.
33.69 Ignitions system.
33.70 Engine life-limited parts.
33.71 Lubrication system.
33.72 Hydraulic actuating systems.
33.73 Power or thrust response.
33.74 Continued rotation.
33.75 Safety analysis.
33.76 Bird ingestion.
33.77 Foreign object ingestion--ice.
33.78 Rain and hail ingestion.
33.79 Fuel burning thrust augmentor.

             Subpart F_Block Tests; Turbine Aircraft Engines

33.81 Applicability.
33.82 General.
33.83 Vibration test.
33.84 Engine overtorque test.
33.85 Calibration tests.
33.87 Endurance test.
33.88 Engine overtemperature test.
33.89 Operation test.
33.90 Initial maintenance inspection test.
33.91 Engine system and component tests.
33.92 Rotor locking tests.
33.93 Teardown inspection.
33.94 Blade containment and rotor unbalance tests.
33.95 Engine-propeller systems tests.
33.96 Engine tests in auxiliary power unit (APU) mode.
33.97 Thrust reversers.
33.99 General conduct of block tests.

        Subpart G_Special Requirements: Turbine Aircraft Engines

33.201 Design and test requirements for Early ETOPS eligibility.

Appendix A to Part 33--Instructions for Continued Airworthiness
Appendix B to Part 33--Certification Standard Atmospheric Concentrations 
          of Rain and Hail
Appendix C to Part 33 [Reserved]
Appendix D to Part 33--Mixed Phase and Ice Crystal Icing Envelope (Deep 
          Convective Clouds)

    Authority: 49 U.S.C. 106(g), 40113, 44701-44702, 44704.

    Effective Date Note: At 79 FR 65536, Nov. 4, 2014, the authority 
citation for part 33 was revised, effective Jan. 5, 2015. For the 
convenience of the user, the revised text is set forth as follows:
    Authority: 49 U.S.C. 106(g), 40113, 44701, 44702, 44704.

    Source: Docket No. 3025, 29 FR 7453, June 10, 1964, unless otherwise 
noted.
    Note: For miscellaneous amendments to cross references in this Part 
33, see Amdt. 33-2, 31 FR 9211, July 6, 1966.



                            Subpart A_General



Sec. 33.1  Applicability.

    (a) This part prescribes airworthiness standards for the issue of 
type certificates and changes to those certificates, for aircraft 
engines.
    (b) Each person who applies under part 21 for such a certificate or 
change must show compliance with the applicable requirements of this 
part and the applicable requirements of part 34 of this chapter.

[Amdt. 33-7, 41 FR 55474, Dec. 20, 1976, as amended by Amdt. 33-14, 55 
FR 32861, Aug. 10, 1990]



Sec. 33.3  General.

    Each applicant must show that the aircraft engine concerned meets 
the applicable requirements of this part.



Sec. 33.4  Instructions for Continued Airworthiness.

    The applicant must prepare Instructions for Continued Airworthiness 
in accordance with appendix A to this part that are acceptable to the 
Administrator. The instructions may be incomplete at type certification 
if a program exists to ensure their completion prior to delivery of the 
first aircraft with the engine installed, or upon issuance of a standard 
certificate of airworthiness for the aircraft with the engine installed, 
whichever occurs later.

[Amdt. 33-9, 45 FR 60181, Sept. 11, 1980]

[[Page 871]]



Sec. 33.5  Instruction manual for installing and operating the engine.

    Each applicant must prepare and make available to the Administrator 
prior to the issuance of the type certificate, and to the owner at the 
time of delivery of the engine, approved instructions for installing and 
operating the engine. The instructions must include at least the 
following:
    (a) Installation instructions. (1) The location of engine mounting 
attachments, the method of attaching the engine to the aircraft, and the 
maximum allowable load for the mounting attachments and related 
structure.
    (2) The location and description of engine connections to be 
attached to accessories, pipes, wires, cables, ducts, and cowling.
    (3) An outline drawing of the engine including overall dimensions.
    (4) A definition of the physical and functional interfaces with the 
aircraft and aircraft equipment, including the propeller when 
applicable.
    (5) Where an engine system relies on components that are not part of 
the engine type design, the interface conditions and reliability 
requirements for those components upon which engine type certification 
is based must be specified in the engine installation instructions 
directly or by reference to appropriate documentation.
    (6) A list of the instruments necessary for control of the engine, 
including the overall limits of accuracy and transient response required 
of such instruments for control of the operation of the engine, must 
also be stated so that the suitability of the instruments as installed 
may be assessed.
    (b) Operation instructions. (1) The operating limitations 
established by the Administrator.
    (2) The power or thrust ratings and procedures for correcting for 
nonstandard atmosphere.
    (3) The recommended procedures, under normal and extreme ambient 
conditions for--
    (i) Starting;
    (ii) Operating on the ground; and
    (iii) Operating during flight.
    (4) For rotorcraft engines having one or more OEI ratings, 
applicants must provide data on engine performance characteristics and 
variability to enable the aircraft manufacturer to establish aircraft 
power assurance procedures.
    (5) A description of the primary and all alternate modes, and any 
back-up system, together with any associated limitations, of the engine 
control system and its interface with the aircraft systems, including 
the propeller when applicable.
    (c) Safety analysis assumptions. The assumptions of the safety 
analysis as described in Sec. 33.75(d) with respect to the reliability 
of safety devices, instrumentation, early warning devices, maintenance 
checks, and similar equipment or procedures that are outside the control 
of the engine manufacturer.

[Amdt. 33-6, 39 FR 35463, Oct. 1, 1974, as amended by Amdt. 33-9, 45 FR 
60181, Sept. 11, 1980; Amdt. 33-24, 47 FR 50867, Sept. 4, 2007; Amdt. 
33-25, 73 FR 48123, Aug. 18, 2008; Amdt. 33-26, 73 FR 48284, Aug. 19, 
2008]



Sec. 33.7  Engine ratings and operating limitations.

    (a) Engine ratings and operating limitations are established by the 
Administrator and included in the engine certificate data sheet 
specified in Sec. 21.41 of this chapter, including ratings and 
limitations based on the operating conditions and information specified 
in this section, as applicable, and any other information found 
necessary for safe operation of the engine.
    (b) For reciprocating engines, ratings and operating limitations are 
established relating to the following:
    (1) Horsepower or torque, r.p.m., manifold pressure, and time at 
critical pressure altitude and sea level pressure altitude for--
    (i) Rated maximum continuous power (relating to unsupercharged 
operation or to operation in each supercharger mode as applicable); and
    (ii) Rated takeoff power (relating to unsupercharged operation or to 
operation in each supercharger mode as applicable).
    (2) Fuel grade or specification.
    (3) Oil grade or specification.
    (4) Temperature of the--
    (i) Cylinder;
    (ii) Oil at the oil inlet; and
    (iii) Turbosupercharger turbine wheel inlet gas.

[[Page 872]]

    (5) Pressure of--
    (i) Fuel at the fuel inlet; and
    (ii) Oil at the main oil gallery.
    (6) Accessory drive torque and overhang moment.
    (7) Component life.
    (8) Turbosupercharger turbine wheel r.p.m.
    (c) For turbine engines, ratings and operating limitations are 
established relating to the following:
    (1) Horsepower, torque, or thrust, r.p.m., gas temperature, and time 
for--
    (i) Rated maximum continuous power or thrust (augmented);
    (ii) Rated maximum continuous power or thrust (unaugmented);
    (iii) Rated takeoff power or thrust (augmented);
    (iv) Rated takeoff power or thrust (unaugmented);
    (v) Rated 30-minute OEI power;
    (vi) Rated 2\1/2\-minute OEI power;
    (vii) Rated continuous OEI power; and
    (viii) Rated 2-minute OEI Power;
    (ix) Rated 30-second OEI power; and
    (x) Auxiliary power unit (APU) mode of operation.
    (2) Fuel designation or specification.
    (3) Oil grade or specification.
    (4) Hydraulic fluid specification.
    (5) Temperature of--
    (i) Oil at a location specified by the applicant;
    (ii) Induction air at the inlet face of a supersonic engine, 
including steady state operation and transient over-temperature and time 
allowed;
    (iii) Hydraulic fluid of a supersonic engine;
    (iv) Fuel at a location specified by the applicant; and
    (v) External surfaces of the engine, if specified by the applicant.
    (6) Pressure of--
    (i) Fuel at the fuel inlet;
    (ii) Oil at a location specified by the applicant;
    (iii) Induction air at the inlet face of a supersonic engine, 
including steady state operation and transient overpressure and time 
allowed; and
    (iv) Hydraulic fluid.
    (7) Accessory drive torque and overhang moment.
    (8) Component life.
    (9) Fuel filtration.
    (10) Oil filtration.
    (11) Bleed air.
    (12) The number of start-stop stress cycles approved for each rotor 
disc and spacer.
    (13) Inlet air distortion at the engine inlet.
    (14) Transient rotor shaft overspeed r.p.m., and number of overspeed 
occurrences.
    (15) Transient gas overtemperature, and number of overtemperature 
occurrences.
    (16) Transient engine overtorque, and number of overtorque 
occurrences.
    (17) Maximum engine overtorque for turbopropeller and turboshaft 
engines incorporating free power turbines.
    (18) For engines to be used in supersonic aircraft, engine rotor 
windmilling rotational r.p.m.
    (d) In determining the engine performance and operating limitations, 
the overall limits of accuracy of the engine control system and of the 
necessary instrumentation as defined in Sec. 33.5(a)(6) must be taken 
into account.

[Amdt. 33-6, 39 FR 35463, Oct. 1, 1974, as amended by Amdt. 33-10, 49 FR 
6850, Feb. 23, 1984; Amdt. 33-11, 51 FR 10346, Mar. 25, 1986; Amdt. 33-
12, 53 FR 34220, Sept. 2, 1988; Amdt. 33-18, 61 FR 31328, June 19, 1996; 
Amdt. 33-26, 73 FR 48284, Aug. 19, 2008; Amdt. 33-30, 74 FR 45310, Sept. 
2, 2009]



Sec. 33.8  Selection of engine power and thrust ratings.

    (a) Requested engine power and thrust ratings must be selected by 
the applicant.
    (b) Each selected rating must be for the lowest power or thrust that 
all engines of the same type may be expected to produce under the 
conditions used to determine that rating.

[Amdt. 33-3, 32 FR 3736, Mar. 4, 1967]



               Subpart B_Design and Construction; General



Sec. 33.11  Applicability.

    This subpart prescribes the general design and construction 
requirements for reciprocating and turbine aircraft engines.

[[Page 873]]



Sec. 33.13  [Reserved]



Sec. 33.15  Materials.

    The suitability and durability of materials used in the engine 
must--
    (a) Be established on the basis of experience or tests; and
    (b) Conform to approved specifications (such as industry or military 
specifications) that ensure their having the strength and other 
properties assumed in the design data.

(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, 
and 1423; sec. 6(c), 49 U.S.C. 1655(c))

[Amdt. 33-8, 42 FR 15047, Mar. 17, 1977, as amended by Amdt. 33-10, 49 
FR 6850, Feb. 23, 1984]



Sec. 33.17  Fire protection.

    (a) The design and construction of the engine and the materials used 
must minimize the probability of the occurrence and spread of fire 
during normal operation and failure conditions, and must minimize the 
effect of such a fire. In addition, the design and construction of 
turbine engines must minimize the probability of the occurrence of an 
internal fire that could result in structural failure or other hazardous 
effects.
    (b) Except as provided in paragraph (c) of this section, each 
external line, fitting, and other component, which contains or conveys 
flammable fluid during normal engine operation, must be fire resistant 
or fireproof, as determined by the Administrator. Components must be 
shielded or located to safeguard against the ignition of leaking 
flammable fluid.
    (c) A tank, which contains flammable fluids and any associated shut-
off means and supports, which are part of and attached to the engine, 
must be fireproof either by construction or by protection unless damage 
by fire will not cause leakage or spillage of a hazardous quantity of 
flammable fluid. For a reciprocating engine having an integral oil sump 
of less than 23.7 liters capacity, the oil sump need not be fireproof or 
enclosed by a fireproof shield.
    (d) An engine component designed, constructed, and installed to act 
as a firewall must be:
    (1) Fireproof;
    (2) Constructed so that no hazardous quantity of air, fluid or flame 
can pass around or through the firewall; and,
    (3) Protected against corrosion;
    (e) In addition to the requirements of paragraphs (a) and (b) of 
this section, engine control system components that are located in a 
designated fire zone must be fire resistant or fireproof, as determined 
by the Administrator.
    (f) Unintentional accumulation of hazardous quantities of flammable 
fluid within the engine must be prevented by draining and venting.
    (g) Any components, modules, or equipment, which are susceptible to 
or are potential sources of static discharges or electrical fault 
currents must be designed and constructed to be properly grounded to the 
engine reference, to minimize the risk of ignition in external areas 
where flammable fluids or vapors could be present.

[Doc. No. FAA-2007-28503, 74 FR 37930, July 30, 2009]



Sec. 33.19  Durability.

    (a) Engine design and construction must minimize the development of 
an unsafe condition of the engine between overhaul periods. The design 
of the compressor and turbine rotor cases must provide for the 
containment of damage from rotor blade failure. Energy levels and 
trajectories of fragments resulting from rotor blade failure that lie 
outside the compressor and turbine rotor cases must be defined.
    (b) Each component of the propeller blade pitch control system which 
is a part of the engine type design must meet the requirements of 
Sec. Sec. 35.21, 35.23, 35.42 and 35.43 of this chapter.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-9, 45 
FR 60181, Sept. 11, 1980; Amdt. 33-10, 49 FR 6851, Feb. 23, 1984; Amdt. 
33-28, 73 FR 63346, Oct. 24, 2008]



Sec. 33.21  Engine cooling.

    Engine design and construction must provide the necessary cooling 
under conditions in which the airplane is expected to operate.

[[Page 874]]



Sec. 33.23  Engine mounting attachments and structure.

    (a) The maximum allowable limit and ultimate loads for engine 
mounting attachments and related engine structure must be specified.
    (b) The engine mounting attachments and related engine structure 
must be able to withstand--
    (1) The specified limit loads without permanent deformation; and
    (2) The specified ultimate loads without failure, but may exhibit 
permanent deformation.

[Amdt. 33-10, 49 FR 6851, Feb. 23, 1984]



Sec. 33.25  Accessory attachments.

    The engine must operate properly with the accessory drive and 
mounting attachments loaded. Each engine accessory drive and mounting 
attachment must include provisions for sealing to prevent contamination 
of, or unacceptable leakage from, the engine interior. A drive and 
mounting attachment requiring lubrication for external drive splines, or 
coupling by engine oil, must include provisions for sealing to prevent 
unacceptable loss of oil and to prevent contamination from sources 
outside the chamber enclosing the drive connection. The design of the 
engine must allow for the examination, adjustment, or removal of each 
accessory required for engine operation.

[Amdt. 33-10, 49 FR 6851, Feb. 23, 1984]



Sec. 33.27  Turbine, compressor, fan, and turbosupercharger rotor
overspeed.

    (a) For each fan, compressor, turbine, and turbosupercharger rotor, 
the applicant must establish by test, analysis, or a combination of 
both, that each rotor will not burst when operated in the engine for 5 
minutes at whichever of the conditions defined in paragraph (b) of this 
section is the most critical with respect to the integrity of such a 
rotor.
    (1) Test rotors used to demonstrate compliance with this section 
that do not have the most adverse combination of material properties and 
dimensional tolerances must be tested at conditions which have been 
adjusted to ensure the minimum specification rotor possesses the 
required overspeed capability. This can be accomplished by increasing 
test speed, temperature, and/or loads.
    (2) When an engine test is being used to demonstrate compliance with 
the overspeed conditions listed in paragraph (b)(3) or (b)(4) of this 
section and the failure of a component or system is sudden and 
transient, it may not be possible to operate the engine for 5 minutes 
after the failure. Under these circumstances, the actual overspeed 
duration is acceptable if the required maximum overspeed is achieved.
    (b) When determining the maximum overspeed condition applicable to 
each rotor in order to comply with paragraphs (a) and (c) of this 
section, the applicant must evaluate the following rotor speeds taking 
into consideration the part's operating temperatures and temperature 
gradients throughout the engine's operating envelope:
    (1) 120 percent of the maximum permissible rotor speed associated 
with any of the engine ratings except one-engine-inoperative (OEI) 
ratings of less than 2\1/2\ minutes.
    (2) 115 percent of the maximum permissible rotor speed associated 
with any OEI ratings of less than 2\1/2\ minutes.
    (3) 105 percent of the highest rotor speed that would result from 
either:
    (i) The failure of the component or system which, in a 
representative installation of the engine, is the most critical with 
respect to overspeed when operating at any rating condition except OEI 
ratings of less than 2\1/2\ minutes, or
    (ii) The failure of any component or system in a representative 
installation of the engine, in combination with any other failure of a 
component or system that would not normally be detected during a routine 
pre-flight check or during normal flight operation, that is the most 
critical with respect to overspeed, except as provided by paragraph (c) 
of this section, when operating at any rating condition except OEI 
ratings of less than 2\1/2\ minutes.
    (4) 100 percent of the highest rotor speed that would result from 
the failure of the component or system which, in a representative 
installation of the engine, is the most critical with respect to 
overspeed when operating at

[[Page 875]]

any OEI rating of less than 2\1/2\ minutes.
    (c) The highest overspeed that results from a complete loss of load 
on a turbine rotor, except as provided by paragraph (f) of this section, 
must be included in the overspeed conditions considered by paragraphs 
(b)(3)(i), (b)(3)(ii), and (b)(4) of this section, regardless of whether 
that overspeed results from a failure within the engine or external to 
the engine. The overspeed resulting from any other single failure must 
be considered when selecting the most limiting overspeed conditions 
applicable to each rotor. Overspeeds resulting from combinations of 
failures must also be considered unless the applicant can show that the 
probability of occurrence is not greater than extremely remote 
(probability range of 10-7 to 10-9 per engine 
flight hour).
    (d) In addition, the applicant must demonstrate that each fan, 
compressor, turbine, and turbosupercharger rotor complies with 
paragraphs (d)(1) and (d)(2) of this section for the maximum overspeed 
achieved when subjected to the conditions specified in paragraphs (b)(3) 
and (b)(4) of this section. The applicant must use the approach in 
paragraph (a) of this section which specifies the required test 
conditions.
    (1) Rotor Growth must not cause the engine to:
    (i) Catch fire,
    (ii) Release high-energy debris through the engine casing or result 
in a hazardous failure of the engine casing,
    (iii) Generate loads greater than those ultimate loads specified in 
Sec. 33.23(a), or
    (iv) Lose the capability of being shut down.
    (2) Following an overspeed event and after continued operation, the 
rotor may not exhibit conditions such as cracking or distortion which 
preclude continued safe operation.
    (e) The design and functioning of engine control systems, 
instruments, and other methods not covered under Sec. 33.28 must ensure 
that the engine operating limitations that affect turbine, compressor, 
fan, and turbosupercharger rotor structural integrity will not be 
exceeded in service.
    (f) Failure of a shaft section may be excluded from consideration in 
determining the highest overspeed that would result from a complete loss 
of load on a turbine rotor if the applicant:
    (1) Identifies the shaft as an engine life-limited-part and complies 
with Sec. 33.70.
    (2) Uses material and design features that are well understood and 
that can be analyzed by well-established and validated stress analysis 
techniques.
    (3) Determines, based on an assessment of the environment 
surrounding the shaft section, that environmental influences are 
unlikely to cause a shaft failure. This assessment must include 
complexity of design, corrosion, wear, vibration, fire, contact with 
adjacent components or structure, overheating, and secondary effects 
from other failures or combination of failures.
    (4) Identifies and declares, in accordance with Sec. 33.5, any 
assumptions regarding the engine installation in making the assessment 
described above in paragraph (f)(3) of this section.
    (5) Assesses, and considers as appropriate, experience with shaft 
sections of similar design.
    (6) Does not exclude the entire shaft.
    (g) If analysis is used to meet the overspeed requirements, then the 
analytical tool must be validated to prior overspeed test results of a 
similar rotor. The tool must be validated for each material. The rotor 
being certified must not exceed the boundaries of the rotors being used 
to validate the analytical tool in terms of geometric shape, operating 
stress, and temperature. Validation includes the ability to accurately 
predict rotor dimensional growth and the burst speed. The predictions 
must also show that the rotor being certified does not have lower burst 
and growth margins than rotors used to validate the tool.

[Doc. No. FAA-2010-0398, Amdt. 33-31, 76 FR 42023, July 18, 2011]



Sec. 33.28  Engine control systems.

    (a) Applicability. These requirements are applicable to any system 
or device that is part of engine type design, that controls, limits, or 
monitors engine operation, and is necessary for the continued 
airworthiness of the engine.

[[Page 876]]

    (b) Validation--(1) Functional aspects. The applicant must 
substantiate by tests, analysis, or a combination thereof, that the 
engine control system performs the intended functions in a manner which:
    (i) Enables selected values of relevant control parameters to be 
maintained and the engine kept within the approved operating limits over 
changing atmospheric conditions in the declared flight envelope;
    (ii) Complies with the operability requirements of Sec. Sec. 33.51, 
33.65 and 33.73, as appropriate, under all likely system inputs and 
allowable engine power or thrust demands, unless it can be demonstrated 
that failure of the control function results in a non-dispatchable 
condition in the intended application;
    (iii) Allows modulation of engine power or thrust with adequate 
sensitivity over the declared range of engine operating conditions; and
    (iv) Does not create unacceptable power or thrust oscillations.
    (2) Environmental limits. The applicant must demonstrate, when 
complying with Sec. Sec. 33.53 or 33.91, that the engine control system 
functionality will not be adversely affected by declared environmental 
conditions, including electromagnetic interference (EMI), High Intensity 
Radiated Fields (HIRF), and lightning. The limits to which the system 
has been qualified must be documented in the engine installation 
instructions.
    (c) Control transitions. (1) The applicant must demonstrate that, 
when fault or failure results in a change from one control mode to 
another, from one channel to another, or from the primary system to the 
back-up system, the change occurs so that:
    (i) The engine does not exceed any of its operating limitations;
    (ii) The engine does not surge, stall, or experience unacceptable 
thrust or power changes or oscillations or other unacceptable 
characteristics; and
    (iii) There is a means to alert the flight crew if the crew is 
required to initiate, respond to, or be aware of the control mode 
change. The means to alert the crew must be described in the engine 
installation instructions, and the crew action must be described in the 
engine operating instructions;
    (2) The magnitude of any change in thrust or power and the 
associated transition time must be identified and described in the 
engine installation instructions and the engine operating instructions.
    (d) Engine control system failures. The applicant must design and 
construct the engine control system so that:
    (1) The rate for Loss of Thrust (or Power) Control (LOTC/LOPC) 
events, consistent with the safety objective associated with the 
intended application can be achieved;
    (2) In the full-up configuration, the system is single fault 
tolerant, as determined by the Administrator, for electrical or 
electronic failures with respect to LOTC/LOPC events;
    (3) Single failures of engine control system components do not 
result in a hazardous engine effect; and
    (4) Foreseeable failures or malfunctions leading to local events in 
the intended aircraft installation, such as fire, overheat, or failures 
leading to damage to engine control system components, do not result in 
a hazardous engine effect due to engine control system failures or 
malfunctions.
    (e) System safety assessment. When complying with this section and 
Sec. 33.75, the applicant must complete a System Safety Assessment for 
the engine control system. This assessment must identify faults or 
failures that result in a change in thrust or power, transmission of 
erroneous data, or an effect on engine operability producing a surge or 
stall together with the predicted frequency of occurrence of these 
faults or failures.
    (f) Protection systems. (1) The design and functioning of engine 
control devices and systems, together with engine instruments and 
operating and maintenance instructions, must provide reasonable 
assurance that those engine operating limitations that affect turbine, 
compressor, fan, and turbosupercharger rotor structural integrity will 
not be exceeded in service.
    (2) When electronic overspeed protection systems are provided, the 
design must include a means for testing, at least once per engine start/
stop cycle, to establish the availability of the protection function. 
The means must be such that a complete test of the system

[[Page 877]]

can be achieved in the minimum number of cycles. If the test is not 
fully automatic, the requirement for a manual test must be contained in 
the engine instructions for operation.
    (3) When overspeed protection is provided through hydromechanical or 
mechanical means, the applicant must demonstrate by test or other 
acceptable means that the overspeed function remains available between 
inspection and maintenance periods.
    (g) Software. The applicant must design, implement, and verify all 
associated software to minimize the existence of errors by using a 
method, approved by the FAA, consistent with the criticality of the 
performed functions.
    (h) Aircraft-supplied data. Single failures leading to loss, 
interruption or corruption of aircraft-supplied data (other than thrust 
or power command signals from the aircraft), or data shared between 
engines must:
    (1) Not result in a hazardous engine effect for any engine; and
    (2) Be detected and accommodated. The accommodation strategy must 
not result in an unacceptable change in thrust or power or an 
unacceptable change in engine operating and starting characteristics. 
The applicant must evaluate and document in the engine installation 
instructions the effects of these failures on engine power or thrust, 
engine operability, and starting characteristics throughout the flight 
envelope.
    (i) Aircraft-supplied electrical power. (1) The applicant must 
design the engine control system so that the loss, malfunction, or 
interruption of electrical power supplied from the aircraft to the 
engine control system will not result in any of the following:
    (i) A hazardous engine effect, or
    (ii) The unacceptable transmission of erroneous data.
    (2) When an engine dedicated power source is required for compliance 
with paragraph (i)(1) of this section, its capacity should provide 
sufficient margin to account for engine operation below idle where the 
engine control system is designed and expected to recover engine 
operation automatically.
    (3) The applicant must identify and declare the need for, and the 
characteristics of, any electrical power supplied from the aircraft to 
the engine control system for starting and operating the engine, 
including transient and steady state voltage limits, in the engine 
instructions for installation.
    (4) Low voltage transients outside the power supply voltage 
limitations declared in paragraph (i)(3) of this section must meet the 
requirements of paragraph (i)(1) of this section. The engine control 
system must be capable of resuming normal operation when aircraft-
supplied power returns to within the declared limits.
    (j) Air pressure signal. The applicant must consider the effects of 
blockage or leakage of the signal lines on the engine control system as 
part of the System Safety Assessment of paragraph (e) of this section 
and must adopt the appropriate design precautions.
    (k) Automatic availability and control of engine power for 30-second 
OEI rating. Rotorcraft engines having a 30-second OEI rating must 
incorporate a means, or a provision for a means, for automatic 
availability and automatic control of the 30-second OEI power within its 
operating limitations.
    (l) Engine shut down means. Means must be provided for shutting down 
the engine rapidly.
    (m) Programmable logic devices. The development of programmable 
logic devices using digital logic or other complex design technologies 
must provide a level of assurance for the encoded logic commensurate 
with the hazard associated with the failure or malfunction of the 
systems in which the devices are located. The applicant must provide 
evidence that the development of these devices has been done by using a 
method, approved by the FAA, that is consistent with the criticality of 
the performed function.

[Amdt. 33-26, 73 FR 48284, Aug. 19, 2008]



Sec. 33.29  Instrument connection.

    (a) Unless it is constructed to prevent its connection to an 
incorrect instrument, each connection provided for powerplant 
instruments required by aircraft airworthiness regulations or necessary 
to insure operation of the engine in compliance with any engine 
limitation must be marked to identify it with its corresponding 
instrument.

[[Page 878]]

    (b) A connection must be provided on each turbojet engine for an 
indicator system to indicate rotor system unbalance.
    (c) Each rotorcraft turbine engine having a 30-second OEI rating and 
a 2-minute OEI rating must have a means or a provision for a means to:
    (1) Alert the pilot when the engine is at the 30-second OEI and the 
2-minute OEI power levels, when the event begins, and when the time 
interval expires;
    (2) Automatically record each usage and duration of power at the 30-
second OEI and 2-minute OEI levels;
    (3) Alert maintenance personnel in a positive manner that the engine 
has been operated at either or both of the 30-second and 2-minute OEI 
power levels, and permit retrieval of the recorded data; and
    (4) Enable routine verification of the proper operation of the above 
means.
    (d) The means, or the provision for a means, of paragraphs (c)(2) 
and (c)(3) of this section must not be capable of being reset in flight.
    (e) The applicant must make provision for the installation of 
instrumentation necessary to ensure operation in compliance with engine 
operating limitations. Where, in presenting the safety analysis, or 
complying with any other requirement, dependence is placed on 
instrumentation that is not otherwise mandatory in the assumed aircraft 
installation, then the applicant must specify this instrumentation in 
the engine installation instructions and declare it mandatory in the 
engine approval documentation.
    (f) As part of the System Safety Assessment of Sec. 33.28(e), the 
applicant must assess the possibility and subsequent effect of incorrect 
fit of instruments, sensors, or connectors. Where necessary, the 
applicant must take design precautions to prevent incorrect 
configuration of the system.
    (g) The sensors, together with associated wiring and signal 
conditioning, must be segregated, electrically and physically, to the 
extent necessary to ensure that the probability of a fault propagating 
from instrumentation and monitoring functions to control functions, or 
vice versa, is consistent with the failure effect of the fault.
    (h) The applicant must provide instrumentation enabling the flight 
crew to monitor the functioning of the turbine cooling system unless 
appropriate inspections are published in the relevant manuals and 
evidence shows that:
    (1) Other existing instrumentation provides adequate warning of 
failure or impending failure;
    (2) Failure of the cooling system would not lead to hazardous engine 
effects before detection; or
    (3) The probability of failure of the cooling system is extremely 
remote.

[Amdt. 33-5, 39 FR 1831, Jan. 15, 1974, as amended by Amdt. 33-6, 39 FR 
35465, Oct. 1, 1974; Amdt. 33-18, 61 FR 31328, June 19, 1996; Amdt. 33-
25, 73 FR 48123, Aug. 18, 2008; Amdt. 33-26, 73 FR 48285, Aug. 19, 2008]



    Subpart C_Design and Construction; Reciprocating Aircraft Engines



Sec. 33.31  Applicability.

    This subpart prescribes additional design and construction 
requirements for reciprocating aircraft engines.



Sec. 33.33  Vibration.

    The engine must be designed and constructed to function throughout 
its normal operating range of crankshaft rotational speeds and engine 
powers without inducing excessive stress in any of the engine parts 
because of vibration and without imparting excessive vibration forces to 
the aircraft structure.



Sec. 33.34  Turbocharger rotors.

    Each turbocharger case must be designed and constructed to be able 
to contain fragments of a compressor or turbine that fails at the 
highest speed that is obtainable with normal speed control devices 
inoperative.

[Amdt. 33-22, 72 FR 50860, Sept. 4, 2007]



Sec. 33.35  Fuel and induction system.

    (a) The fuel system of the engine must be designed and constructed 
to supply an appropriate mixture of fuel to the cylinders throughout the 
complete operating range of the engine under all flight and atmospheric 
conditions.

[[Page 879]]

    (b) The intake passages of the engine through which air or fuel in 
combination with air passes for combustion purposes must be designed and 
constructed to minimize the danger of ice accretion in those passages. 
The engine must be designed and constructed to permit the use of a means 
for ice prevention.
    (c) The type and degree of fuel filtering necessary for protection 
of the engine fuel system against foreign particles in the fuel must be 
specified. The applicant must show that foreign particles passing 
through the prescribed filtering means will not critically impair engine 
fuel system functioning.
    (d) Each passage in the induction system that conducts a mixture of 
fuel and air must be self-draining, to prevent a liquid lock in the 
cylinders, in all attitudes that the applicant establishes as those the 
engine can have when the aircraft in which it is installed is in the 
static ground attitude.
    (e) If provided as part of the engine, the applicant must show for 
each fluid injection (other than fuel) system and its controls that the 
flow of the injected fluid is adequately controlled.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-10, 49 
FR 6851, Feb. 23, 1984]



Sec. 33.37  Ignition system.

    Each spark ignition engine must have a dual ignition system with at 
least two spark plugs for each cylinder and two separate electric 
circuits with separate sources of electrical energy, or have an ignition 
system of equivalent in-flight reliability.



Sec. 33.39  Lubrication system.

    (a) The lubrication system of the engine must be designed and 
constructed so that it will function properly in all flight attitudes 
and atmospheric conditions in which the airplane is expected to operate. 
In wet sump engines, this requirement must be met when only one-half of 
the maximum lubricant supply is in the engine.
    (b) The lubrication system of the engine must be designed and 
constructed to allow installing a means of cooling the lubricant.
    (c) The crankcase must be vented to the atmosphere to preclude 
leakage of oil from excessive pressure in the crankcase.



          Subpart D_Block Tests; Reciprocating Aircraft Engines



Sec. 33.41  Applicability.

    This subpart prescribes the block tests and inspections for 
reciprocating aircraft engines.



Sec. 33.42  General.

    Before each endurance test required by this subpart, the adjustment 
setting and functioning characteristic of each component having an 
adjustment setting and a functioning characteristic that can be 
established independent of installation on the engine must be 
established and recorded.

[Amdt. 33-6, 39 FR 35465, Oct. 1, 1974]



Sec. 33.43  Vibration test.

    (a) Each engine must undergo a vibration survey to establish the 
torsional and bending vibration characteristics of the crankshaft and 
the propeller shaft or other output shaft, over the range of crankshaft 
speed and engine power, under steady state and transient conditions, 
from idling speed to either 110 percent of the desired maximum 
continuous speed rating or 103 percent of the maximum desired takeoff 
speed rating, whichever is higher. The survey must be conducted using, 
for airplane engines, the same configuration of the propeller type which 
is used for the endurance test, and using, for other engines, the same 
configuration of the loading device type which is used for the endurance 
test.
    (b) The torsional and bending vibration stresses of the crankshaft 
and the propeller shaft or other output shaft may not exceed the 
endurance limit stress of the material from which the shaft is made. If 
the maximum stress in the shaft cannot be shown to be below the 
endurance limit by measurement, the vibration frequency and amplitude 
must be measured. The peak amplitude must be shown to produce a stress 
below the endurance limit; if not, the engine must be run at the 
condition producing the peak amplitude until, for steel shafts, 10 
million stress

[[Page 880]]

reversals have been sustained without fatigue failure and, for other 
shafts, until it is shown that fatigue will not occur within the 
endurance limit stress of the material.
    (c) Each accessory drive and mounting attachment must be loaded, 
with the loads imposed by each accessory used only for an aircraft 
service being the limit load specified by the applicant for the drive or 
attachment point.
    (d) The vibration survey described in paragraph (a) of this section 
must be repeated with that cylinder not firing which has the most 
adverse vibration effect, in order to establish the conditions under 
which the engine can be operated safely in that abnormal state. However, 
for this vibration survey, the engine speed range need only extend from 
idle to the maximum desired takeoff speed, and compliance with paragraph 
(b) of this section need not be shown.

[Amdt. 33-6, 39 FR 35465, Oct. 1, 1974, as amended by Amdt. 33-10, 49 FR 
6851, Feb. 23, 1984]



Sec. 33.45  Calibration tests.

    (a) Each engine must be subjected to the calibration tests necessary 
to establish its power characteristics and the conditions for the 
endurance test specified in Sec. 33.49. The results of the power 
characteristics calibration tests form the basis for establishing the 
characteristics of the engine over its entire operating range of 
crankshaft rotational speeds, manifold pressures, fuel/air mixture 
settings, and altitudes. Power ratings are based upon standard 
atmospheric conditions with only those accessories installed which are 
essential for engine functioning.
    (b) A power check at sea level conditions must be accomplished on 
the endurance test engine after the endurance test. Any change in power 
characteristics which occurs during the endurance test must be 
determined. Measurements taken during the final portion of the endurance 
test may be used in showing compliance with the requirements of this 
paragraph.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-6, 39 
FR 35465, Oct. 1, 1974]



Sec. 33.47  Detonation test.

    Each engine must be tested to establish that the engine can function 
without detonation throughout its range of intended conditions of 
operation.



Sec. 33.49  Endurance test.

    (a) General. Each engine must be subjected to an endurance test that 
includes a total of 150 hours of operation (except as provided in 
paragraph (e)(1)(iii) of this section) and, depending upon the type and 
contemplated use of the engine, consists of one of the series of runs 
specified in paragraphs (b) through (e) of this section, as applicable. 
The runs must be made in the order found appropriate by the 
Administrator for the particular engine being tested. During the 
endurance test the engine power and the crankshaft rotational speed must 
be kept within 3 percent of the rated values. 
During the runs at rated takeoff power and for at least 35 hours at 
rated maximum continuous power, one cylinder must be operated at not 
less than the limiting temperature, the other cylinders must be operated 
at a temperature not lower than 50 degrees F. below the limiting 
temperature, and the oil inlet temperature must be maintained within 
10 degrees F. of the limiting temperature. An 
engine that is equipped with a propeller shaft must be fitted for the 
endurance test with a propeller that thrust-loads the engine to the 
maximum thrust which the engine is designed to resist at each applicable 
operating condition specified in this section. Each accessory drive and 
mounting attachment must be loaded. During operation at rated takeoff 
power and rated maximum continuous power, the load imposed by each 
accessory used only for an aircraft service must be the limit load 
specified by the applicant for the engine drive or attachment point.
    (b) Unsupercharged engines and engines incorporating a gear-driven 
single-speed supercharger. For engines not incorporating a supercharger 
and for engines incorporating a gear-driven single-speed supercharger 
the applicant must conduct the following runs:
    (1) A 30-hour run consisting of alternate periods of 5 minutes at 
rated takeoff power with takeoff speed, and 5

[[Page 881]]

minutes at maximum best economy cruising power or maximum recommended 
cruising power.
    (2) A 20-hour run consisting of alternate periods of 1\1/2\ hours at 
rated maximum continuous power with maximum continuous speed, and \1/2\ 
hour at 75 percent rated maximum continuous power and 91 percent maximum 
continuous speed.
    (3) A 20-hour run consisting of alternate periods of 1\1/2\ hours at 
rated maximum continuous power with maximum continuous speed, and \1/2\ 
hour at 70 percent rated maximum continuous power and 89 percent maximum 
continuous speed.
    (4) A 20-hour run consisting of alternate periods of 1\1/2\ hours at 
rated maximum continuous power with maximum continuous speed, and \1/2\ 
hour at 65 percent rated maximum continuous power and 87 percent maximum 
continuous speed.
    (5) A 20-hour run consisting of alternate periods of 1\1/2\ hours at 
rated maximum continuous power with maximum continuous speed, and \1/2\ 
hour at 60 percent rated maximum continuous power and 84.5 percent 
maximum continuous speed.
    (6) A 20-hour run consisting of alternate periods of 1\1/2\ hours at 
rated maximum continuous power with maximum continuous speed, and \1/2\ 
hour at 50 percent rated maximum continuous power and 79.5 percent 
maximum continuous speed.
    (7) A 20-hour run consisting of alternate periods of 2\1/2\ hours at 
rated maximum continuous power with maximum continuous speed, and 2\1/2\ 
hours at maximum best economy cruising power or at maximum recommended 
cruising power.
    (c) Engines incorporating a gear-driven two-speed supercharger. For 
engines incorporating a gear-driven two-speed supercharger the applicant 
must conduct the following runs:
    (1) A 30-hour run consisting of alternate periods in the lower gear 
ratio of 5 minutes at rated takeoff power with takeoff speed, and 5 
minutes at maximum best economy cruising power or at maximum recommended 
cruising power. If a takeoff power rating is desired in the higher gear 
ratio, 15 hours of the 30-hour run must be made in the higher gear ratio 
in alternate periods of 5 minutes at the observed horsepower obtainable 
with the takeoff critical altitude manifold pressure and takeoff speed, 
and 5 minutes at 70 percent high ratio rated maximum continuous power 
and 89 percent high ratio maximum continuous speed.
    (2) A 15-hour run consisting of alternate periods in the lower gear 
ratio of 1 hour at rated maximum continuous power with maximum 
continuous speed, and \1/2\ hour at 75 percent rated maximum continuous 
power and 91 percent maximum continuous speed.
    (3) A 15-hour run consisting of alternate periods in the lower gear 
ratio of 1 hour at rated maximum continuous power with maximum 
continuous speed, and \1/2\ hour at 70 percent rated maximum continuous 
power and 89 percent maximum continuous speed.
    (4) A 30-hour run in the higher gear ratio at rated maximum 
continuous power with maximum continuous speed.
    (5) A 5-hour run consisting of alternate periods of 5 minutes in 
each of the supercharger gear ratios. The first 5 minutes of the test 
must be made at maximum continuous speed in the higher gear ratio and 
the observed horsepower obtainable with 90 percent of maximum continuous 
manifold pressure in the higher gear ratio under sea level conditions. 
The condition for operation for the alternate 5 minutes in the lower 
gear ratio must be that obtained by shifting to the lower gear ratio at 
constant speed.
    (6) A 10-hour run consisting of alternate periods in the lower gear 
ratio of 1 hour at rated maximum continuous power with maximum 
continuous speed, and 1 hour at 65 percent rated maximum continuous 
power and 87 percent maximum continuous speed.
    (7) A 10-hour run consisting of alternate periods in the lower gear 
ratio of 1 hour at rated maximum continuous power with maximum 
continuous speed, and 1 hour at 60 percent rated maximum continuous 
power and 84.5 percent maximum continuous speed.
    (8) A 10-hour run consisting of alternate periods in the lower gear 
ratio of 1 hour at rated maximum continuous power with maximum 
continuous speed, and 1 hour at 50 percent rated

[[Page 882]]

maximum continuous power and 79.5 percent maximum continuous speed.
    (9) A 20-hour run consisting of alternate periods in the lower gear 
ratio of 2 hours at rated maximum continuous power with maximum 
continuous speed, and 2 hours at maximum best economy cruising power and 
speed or at maximum recommended cruising power.
    (10) A 5-hour run in the lower gear ratio at maximum best economy 
cruising power and speed or at maximum recommended cruising power and 
speed.


Where simulated altitude test equipment is not available when operating 
in the higher gear ratio, the runs may be made at the observed 
horsepower obtained with the critical altitude manifold pressure or 
specified percentages thereof, and the fuel-air mixtures may be adjusted 
to be rich enough to suppress detonation.
    (d) Helicopter engines. To be eligible for use on a helicopter each 
engine must either comply with paragraphs (a) through (j) of Sec. 
29.923 of this chapter, or must undergo the following series of runs:
    (1) A 35-hour run consisting of alternate periods of 30 minutes each 
at rated takeoff power with takeoff speed, and at rated maximum 
continuous power with maximum continuous speed.
    (2) A 25-hour run consisting of alternate periods of 2\1/2\ hours 
each at rated maximum continuous power with maximum continuous speed, 
and at 70 percent rated maximum continuous power with maximum continuous 
speed.
    (3) A 25-hour run consisting of alternate periods of 2\1/2\ hours 
each at rated maximum continuous power with maximum continuous speed, 
and at 70 percent rated maximum continuous power with 80 to 90 percent 
maximum continuous speed.
    (4) A 25-hour run consisting of alternate periods of 2\1/2\ hours 
each at 30 percent rated maximum continuous power with takeoff speed, 
and at 30 percent rated maximum continuous power with 80 to 90 percent 
maximum continuous speed.
    (5) A 25-hour run consisting of alternate periods of 2\1/2\ hours 
each at 80 percent rated maximum continuous power with takeoff speed, 
and at either rated maximum continuous power with 110 percent maximum 
continuous speed or at rated takeoff power with 103 percent takeoff 
speed, whichever results in the greater speed.
    (6) A 15-hour run at 105 percent rated maximum continuous power with 
105 percent maximum continuous speed or at full throttle and 
corresponding speed at standard sea level carburetor entrance pressure, 
if 105 percent of the rated maximum continuous power is not exceeded.
    (e) Turbosupercharged engines. For engines incorporating a 
turbosupercharger the following apply except that altitude testing may 
be simulated provided the applicant shows that the engine and 
supercharger are being subjected to mechanical loads and operating 
temperatures no less severe than if run at actual altitude conditions:
    (1) For engines used in airplanes the applicant must conduct the 
runs specified in paragraph (b) of this section, except--
    (i) The entire run specified in paragraph (b)(1) of this section 
must be made at sea level altitude pressure;
    (ii) The portions of the runs specified in paragraphs (b)(2) through 
(7) of this section at rated maximum continuous power must be made at 
critical altitude pressure, and the portions of the runs at other power 
must be made at 8,000 feet altitude pressure; and
    (iii) The turbosupercharger used during the 150-hour endurance test 
must be run on the bench for an additional 50 hours at the limiting 
turbine wheel inlet gas temperature and rotational speed for rated 
maximum continuous power operation unless the limiting temperature and 
speed are maintained during 50 hours of the rated maximum continuous 
power operation.
    (2) For engines used in helicopters the applicant must conduct the 
runs specified in paragraph (d) of this section, except--
    (i) The entire run specified in paragraph (d)(1) of this section 
must be made at critical altitude pressure;
    (ii) The portions of the runs specified in paragraphs (d)(2) and (3) 
of this section at rated maximum continuous

[[Page 883]]

power must be made at critical altitude pressure and the portions of the 
runs at other power must be made at 8,000 feet altitude pressure;
    (iii) The entire run specified in paragraph (d)(4) of this section 
must be made at 8,000 feet altitude pressure;
    (iv) The portion of the runs specified in paragraph (d)(5) of this 
section at 80 percent of rated maximum continuous power must be made at 
8,000 feet altitude pressure and the portions of the runs at other power 
must be made at critical altitude pressure;
    (v) The entire run specified in paragraph (d)(6) of this section 
must be made at critical altitude pressure; and
    (vi) The turbosupercharger used during the endurance test must be 
run on the bench for 50 hours at the limiting turbine wheel inlet gas 
temperature and rotational speed for rated maximum continuous power 
operation unless the limiting temperature and speed are maintained 
during 50 hours of the rated maximum continuous power operation.

[Amdt. 33-3, 32 FR 3736, Mar. 4, 1967, as amended by Amdt. 33-6, 39 FR 
35465, Oct. 1, 1974; Amdt. 33-10, 49 FR 6851, Feb. 23, 1984]



Sec. 33.51  Operation test.

    The operation test must include the testing found necessary by the 
Administrator to demonstrate backfire characteristics, starting, idling, 
acceleration, overspeeding, functioning of propeller and ignition, and 
any other operational characteristic of the engine. If the engine 
incorporates a multispeed supercharger drive, the design and 
construction must allow the supercharger to be shifted from operation at 
the lower speed ratio to the higher and the power appropriate to the 
manifold pressure and speed settings for rated maximum continuous power 
at the higher supercharger speed ratio must be obtainable within five 
seconds.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-3, 32 
FR 3737, Mar. 4, 1967]



Sec. 33.53  Engine system and component tests.

    (a) For those systems and components that cannot be adequately 
substantiated in accordance with endurance testing of Sec. 33.49, the 
applicant must conduct additional tests to demonstrate that systems or 
components are able to perform the intended functions in all declared 
environmental and operating conditions.
    (b) Temperature limits must be established for each component that 
requires temperature controlling provisions in the aircraft installation 
to assure satisfactory functioning, reliability, and durability.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-26, 73 
FR 48285, Aug. 19, 2008]



Sec. 33.55  Teardown inspection.

    After completing the endurance test--
    (a) Each engine must be completely disassembled;
    (b) Each component having an adjustment setting and a functioning 
characteristic that can be established independent of installation on 
the engine must retain each setting and functioning characteristic 
within the limits that were established and recorded at the beginning of 
the test; and
    (c) Each engine component must conform to the type design and be 
eligible for incorporation into an engine for continued operation, in 
accordance with information submitted in compliance with Sec. 33.4.

[Amdt. 33-6, 39 FR 35466, Oct. 1, 1974, as amended by Amdt. 33-9, 45 FR 
60181, Sept. 11, 1980]



Sec. 33.57  General conduct of block tests.

    (a) The applicant may, in conducting the block tests, use separate 
engines of identical design and construction in the vibration, 
calibration, detonation, endurance, and operation tests, except that, if 
a separate engine is used for the endurance test it must be subjected to 
a calibration check before starting the endurance test.
    (b) The applicant may service and make minor repairs to the engine 
during the block tests in accordance with the service and maintenance 
instructions submitted in compliance with Sec. 33.4. If the frequency 
of the service is excessive, or the number of stops due to engine 
malfunction is excessive, or a major repair, or replacement of a part is 
found necessary during the block tests or as the result of findings from

[[Page 884]]

the teardown inspection, the engine or its parts may be subjected to any 
additional test the Administrator finds necessary.
    (c) Each applicant must furnish all testing facilities, including 
equipment and competent personnel, to conduct the block tests.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-6, 39 
FR 35466, Oct. 1, 1974; Amdt. 33-9, 45 FR 60181, Sept. 11, 1980]



       Subpart E_Design and Construction; Turbine Aircraft Engines



Sec. 33.61  Applicability.

    This subpart prescribes additional design and construction 
requirements for turbine aircraft engines.



Sec. 33.62  Stress analysis.

    A stress analysis must be performed on each turbine engine showing 
the design safety margin of each turbine engine rotor, spacer, and rotor 
shaft.

[Amdt. 33-6, 39 FR 35466, Oct. 1, 1974]



Sec. 33.63  Vibration.

    Each engine must be designed and constructed to function throughout 
its declared flight envelope and operating range of rotational speeds 
and power/thrust, without inducing excessive stress in any engine part 
because of vibration and without imparting excessive vibration forces to 
the aircraft structure.

[Doc. No. 28107, 61 FR 28433, June 4, 1996]



Sec. 33.64  Pressurized engine static parts.

    (a) Strength. The applicant must establish by test, validated 
analysis, or a combination of both, that all static parts subject to 
significant gas or liquid pressure loads for a stabilized period of one 
minute will not:
    (1) Exhibit permanent distortion beyond serviceable limits or 
exhibit leakage that could create a hazardous condition when subjected 
to the greater of the following pressures:
    (i) 1.1 times the maximum working pressure;
    (ii) 1.33 times the normal working pressure; or
    (iii) 35 kPa (5 p.s.i.) above the normal working pressure.
    (2) Exhibit fracture or burst when subjected to the greater of the 
following pressures:
    (i) 1.15 times the maximum possible pressure;
    (ii) 1.5 times the maximum working pressure; or
    (iii) 35 kPa (5 p.s.i.) above the maximum possible pressure.
    (b) Compliance with this section must take into account:
    (1) The operating temperature of the part;
    (2) Any other significant static loads in addition to pressure 
loads;
    (3) Minimum properties representative of both the material and the 
processes used in the construction of the part; and
    (4) Any adverse geometry conditions allowed by the type design.

[Amdt. 33-27; 73 FR 55437, Sept. 25, 2008; Amdt. 33-27, 73 FR 57235, 
Oct. 2, 2008]



Sec. 33.65  Surge and stall characteristics.

    When the engine is operated in accordance with operating 
instructions required by Sec. 33.5(b), starting, a change of power or 
thrust, power or thrust augmentation, limiting inlet air distortion, or 
inlet air temperature may not cause surge or stall to the extent that 
flameout, structural failure, overtemperature, or failure of the engine 
to recover power or thrust will occur at any point in the operating 
envelope.

[Amdt. 33-6, 39 FR 35466, Oct. 1, 1974]



Sec. 33.66  Bleed air system.

    The engine must supply bleed air without adverse effect on the 
engine, excluding reduced thrust or power output, at all conditions up 
to the discharge flow conditions established as a limitation under Sec. 
33.7(c)(11). If bleed air used for engine anti-icing can be controlled, 
provision must be made for a means to indicate the functioning of the 
engine ice protection system.

[Amdt. 33-10, 49 FR 6851, Feb. 23, 1984]



Sec. 33.67  Fuel system.

    (a) With fuel supplied to the engine at the flow and pressure 
specified by the applicant, the engine must function properly under each 
operating condition required by this part. Each fuel control adjusting 
means that may

[[Page 885]]

not be manipulated while the fuel control device is mounted on the 
engine must be secured by a locking device and sealed, or otherwise be 
inaccessible. All other fuel control adjusting means must be accessible 
and marked to indicate the function of the adjustment unless the 
function is obvious.
    (b) There must be a fuel strainer or filter between the engine fuel 
inlet opening and the inlet of either the fuel metering device or the 
engine-driven positive displacement pump whichever is nearer the engine 
fuel inlet. In addition, the following provisions apply to each strainer 
or filter required by this paragraph (b):
    (1) It must be accessible for draining and cleaning and must 
incorporate a screen or element that is easily removable.
    (2) It must have a sediment trap and drain except that it need not 
have a drain if the strainer or filter is easily removable for drain 
purposes.
    (3) It must be mounted so that its weight is not supported by the 
connecting lines or by the inlet or outlet connections of the strainer 
or filter, unless adequate strength margins under all loading conditions 
are provided in the lines and connections.
    (4) It must have the type and degree of fuel filtering specified as 
necessary for protection of the engine fuel system against foreign 
particles in the fuel. The applicant must show:
    (i) That foreign particles passing through the specified filtering 
means do not impair the engine fuel system functioning; and
    (ii) That the fuel system is capable of sustained operation 
throughout its flow and pressure range with the fuel initially saturated 
with water at 80 [deg]F (27 [deg]C) and having 0.025 fluid ounces per 
gallon (0.20 milliliters per liter) of free water added and cooled to 
the most critical condition for icing likely to be encountered in 
operation. However, this requirement may be met by demonstrating the 
effectiveness of specified approved fuel anti-icing additives, or that 
the fuel system incorporates a fuel heater which maintains the fuel 
temperature at the fuel strainer or fuel inlet above 32 [deg]F (0 
[deg]C) under the most critical conditions.
    (5) The applicant must demonstrate that the filtering means has the 
capacity (with respect to engine operating limitations) to ensure that 
the engine will continue to operate within approved limits, with fuel 
contaminated to the maximum degree of particle size and density likely 
to be encountered in service. Operation under these conditions must be 
demonstrated for a period acceptable to the Administrator, beginning 
when indication of impending filter blockage is first given by either:
    (i) Existing engine instrumentation; or
    (ii) Additional means incorporated into the engine fuel system.
    (6) Any strainer or filter bypass must be designed and constructed 
so that the release of collected contaminants is minimized by 
appropriate location of the bypass to ensure that collected contaminants 
are not in the bypass flow path.
    (c) If provided as part of the engine, the applicant must show for 
each fluid injection (other than fuel) system and its controls that the 
flow of the injected fluid is adequately controlled.

[Amdt. 33-6, 39 FR 35466, Oct. 1, 1974, as amended by Amdt. 33-10, 49 FR 
6851, Feb. 23, 1984; Amdt. 33-18, 61 FR 31328, June 19, 1996; Amdt. 33-
25, 73 FR 48123, Aug. 18, 2008; Amdt. 33-26, 73 FR 48285, Aug. 19, 2008]



Sec. 33.68  Induction system icing.

    Each engine, with all icing protection systems operating, must--
    (a) Operate throughout its flight power range (including idling) 
without the accumulation of ice on the engine components that adversely 
affects engine operation or that causes a serious loss of power or 
thrust in continuous maximum and intermittent maximum icing conditions 
as defined in appendix C of Part 25 of this chapter; and
    (b) Idle for 30 minutes on the ground, with the available air bleed 
for icing protection at its critical condition, without adverse effect, 
in an atmosphere that is at a temperature between 15[deg] and 30 [deg]F 
(between -9[deg] and -1 [deg]C) and has a liquid water content not less 
than 0.3 grams per cubic meter in the form of drops having a mean 
effective diameter not less than 20 microns, followed by a momentary 
operation at

[[Page 886]]

takeoff power or thrust. During the 30 minutes of idle operation the 
engine may be run up periodically to a moderate power or thrust setting 
in a manner acceptable to the Administrator.

[Amdt. 33-6, 39 FR 35466, Oct. 1, 1974, as amended by Amdt. 33-10, 49 FR 
6852, Feb. 23, 1984]

    Effective Date Note: At 79 FR 65536, Nov. 4, 2014, Sec. 33.68 was 
revised, effective Jan. 5, 2015. For the convenience of the user, the 
revised text is set forth as follows:



Sec. 33.68  Induction system icing.

    Each engine, with all icing protection systems operating, must:
    (a) Operate throughout its flight power range, including the minimum 
descent idle rotor speeds achievable in flight, in the icing conditions 
defined for turbojet, turbofan, and turboprop engines in Appendices C 
and O of part 25 of this chapter, and Appendix D of this part, and for 
turboshaft engines in Appendix C of part 29 of this chapter, without the 
accumulation of ice on the engine components that:
    (1) Adversely affects engine operation or that causes an 
unacceptable permanent loss of power or thrust or unacceptable increase 
in engine operating temperature; or
    (2) Results in unacceptable temporary power loss or engine damage; 
or
    (3) Causes a stall, surge, or flameout or loss of engine 
controllability. The applicant must account for in-flight ram effects in 
any critical point analysis or test demonstration of these flight 
conditions.
    (b) Operate throughout its flight power range, including minimum 
descent idle rotor speeds achievable in flight, in the icing conditions 
defined for turbojet, turbofan, and turboprop engines in Appendices C 
and O of part 25 of this chapter, and for turboshaft engines in Appendix 
C of part 29 of this chapter. In addition:
    (1) It must be shown through Critical Point Analysis (CPA) that the 
complete ice envelope has been analyzed, and that the most critical 
points must be demonstrated by engine test, analysis, or a combination 
of the two to operate acceptably. Extended flight in critical flight 
conditions such as hold, descent, approach, climb, and cruise, must be 
addressed, for the ice conditions defined in these appendices.
    (2) It must be shown by engine test, analysis, or a combination of 
the two that the engine can operate acceptably for the following 
durations:
    (i) At engine powers that can sustain level flight: A duration that 
achieves repetitive, stabilized operation for turbojet, turbofan, and 
turboprop engines in the icing conditions defined in Appendices C and O 
of part 25 of this chapter, and for turboshaft engines in the icing 
conditions defined in Appendix C of part 29 of this chapter.
    (ii) At engine power below that which can sustain level flight:
    (A) Demonstration in altitude flight simulation test facility: A 
duration of 10 minutes consistent with a simulated flight descent of 
10,000 ft (3 km) in altitude while operating in Continuous Maximum icing 
conditions defined in Appendix C of part 25 of this chapter for 
turbojet, turbofan, and turboprop engines, and for turboshaft engines in 
the icing conditions defined in Appendix C of part 29 of this chapter, 
plus 40 percent liquid water content margin, at the critical level of 
airspeed and air temperature; or
    (B) Demonstration in ground test facility: A duration of 3 cycles of 
alternating icing exposure corresponding to the liquid water content 
levels and standard cloud lengths starting in Intermittent Maximum and 
then in Continuous Maximum icing conditions defined in Appendix C of 
part 25 of this chapter for turbojet, turbofan, and turboprop engines, 
and for turboshaft engines in the icing conditions defined in Appendix C 
of part 29 of this chapter, at the critical level of air temperature.
    (c) In addition to complying with paragraph (b) of this section, the 
following conditions shown in Table 1 of this section unless replaced by 
similar CPA test conditions that are more critical or produce an 
equivalent level of severity, must be demonstrated by an engine test:

                         Table 1--Conditions That Must Be Demonstrated by an Engine Test
----------------------------------------------------------------------------------------------------------------
                                                    Supercooled
                                   Total air           water
          Condition               temperature     concentrations   Median volume drop diameter      Duration
                                                     (minimum)
----------------------------------------------------------------------------------------------------------------
1. Glaze ice conditions......  21 to 25 [deg]F   2 g/m\3\........  25 to 35 microns...........  (a) 10-minutes
                                (-6 to -4                                                        for power below
                                [deg]C).                                                         sustainable
                                                                                                 level flight
                                                                                                 (idle descent).
                                                                                                (b) Must show
                                                                                                 repetitive,
                                                                                                 stabilized
                                                                                                 operation for
                                                                                                 higher powers
                                                                                                 (50%, 75%,
                                                                                                 100%MC).

[[Page 887]]

 
2. Rime ice conditions.......  -10 to 0 [deg]F   1 g/m\3\........  15 to 25 microns...........  (a) 10-minutes
                                (-23 to -18                                                      for power below
                                [deg]C).                                                         sustainable
                                                                                                 level flight
                                                                                                 (idle descent).
                                                                                                (b) Must show
                                                                                                 repetitive,
                                                                                                 stabilized
                                                                                                 operation for
                                                                                                 higher powers
                                                                                                 (50%, 75%,
                                                                                                 100%MC).
3. Glaze ice holding           Turbojet and      Alternating       20 to 30 microns...........  Must show
 conditions.                    Turbofan, only:   cycle: First                                   repetitive,
(Turbojet, turbofan, and        10 to 18 [deg]F   1.7 g/m\3\ (1                                  stabilized
 turboprop only).               (-12 to -8        minute), Then                                  operation (or
                                [deg]C).          0.3 g/m\3\ (6                                  45 minutes
                                                  minute).                                       max).
                               Turboprop, only:  ................
                                2 to 10 [deg]F
                                (-17 to -12
                                [deg]C).
4. Rime ice holding            Turbojet and      0.25 g/m\3\.....  20 to 30 microns...........  Must show
 conditions.                    Turbofan, only:                                                  repetitive,
(Turbojet, turbofan, and        -10 to 0 [deg]F                                                  stabilized
 turboprop only).               (-23 to -18                                                      operation (or
                                [deg]C).                                                         45 minutes
                                                                                                 max).
                               Turboprop, only:  ................  ...........................  ................
                                2 to 10 [deg]F
                                (-17 to -12
                                [deg]C).
----------------------------------------------------------------------------------------------------------------

    (d) Operate at ground idle speed for a minimum of 30 minutes at each 
of the following icing conditions shown in Table 2 of this section with 
the available air bleed for icing protection at its critical condition, 
without adverse effect, followed by acceleration to takeoff power or 
thrust. During the idle operation, the engine may be run up periodically 
to a moderate power or thrust setting in a manner acceptable to the 
Administrator. Analysis may be used to show ambient temperatures below 
the tested temperature are less critical. The applicant must document 
any demonstrated run ups and minimum ambient temperature capability in 
the engine operating manual as mandatory in icing conditions. The 
applicant must demonstrate, with consideration of expected airport 
elevations, the following:

                          Table 2--Demonstration Methods for Specific Icing Conditions
----------------------------------------------------------------------------------------------------------------
                                                    Supercooled
                                   Total air           water         Mean effective particle
          Condition               temperature     concentrations             diameter             Demonstration
                                                     (minimum)
----------------------------------------------------------------------------------------------------------------
1. Rime ice condition........  0 to 15 [deg]F (- Liquid--0.3 g/    15-25 microns..............  By engine test.
                                18 to -9          m\3\.
                                [deg]C).
2. Glaze ice condition.......  20 to 30 [deg]F   Liquid--0.3 g/    15-25 microns..............  By engine test.
                                (-7 to -1         m\3\.
                                [deg]C).
3. Snow ice condition........  26 to 32 [deg]F   Ice--0.9 g/m\3\.  100 microns................  By test,
                                (-3 to 0                           (minimum)..................   analysis or
                                [deg]C).                                                         combination of
                                                                                                 the two.
4. Large drop glaze ice        15 to 30 [deg]F   Liquid--0.3 g/    100 microns (minimum)......  By test,
 condition (Turbojet,           (-9 to -1         m\3\.                                          analysis or
 turbofan, and turboprop        [deg]C).                                                         combination of
 only).                                                                                          the two.
----------------------------------------------------------------------------------------------------------------

    (e) Demonstrate by test, analysis, or combination of the two, 
acceptable operation for turbojet, turbofan, and turboprop engines in 
mixed phase and ice crystal icing conditions throughout Appendix D of 
this part, icing envelope throughout its flight power range, including 
minimum descent idling speeds.



Sec. 33.69  Ignitions system.

    Each engine must be equipped with an ignition system for starting 
the engine on the ground and in flight. An electric ignition system must 
have at least two igniters and two separate secondary electric circuits, 
except that

[[Page 888]]

only one igniter is required for fuel burning augmentation systems.

[Amdt. 33-6, 39 FR 35466, Oct. 1, 1974]



Sec. 33.70  Engine life-limited parts.

    By a procedure approved by the FAA, operating limitations must be 
established which specify the maximum allowable number of flight cycles 
for each engine life-limited part. Engine life-limited parts are rotor 
and major static structural parts whose primary failure is likely to 
result in a hazardous engine effect. Typically, engine life-limited 
parts include, but are not limited to disks, spacers, hubs, shafts, 
high-pressure casings, and non-redundant mount components. For the 
purposes of this section, a hazardous engine effect is any of the 
conditions listed in Sec. 33.75 of this part. The applicant will 
establish the integrity of each engine life-limited part by:
    (a) An engineering plan that contains the steps required to ensure 
each engine life-limited part is withdrawn from service at an approved 
life before hazardous engine effects can occur. These steps include 
validated analysis, test, or service experience which ensures that the 
combination of loads, material properties, environmental influences and 
operating conditions, including the effects of other engine parts 
influencing these parameters, are sufficiently well known and 
predictable so that the operating limitations can be established and 
maintained for each engine life-limited part. Applicants must perform 
appropriate damage tolerance assessments to address the potential for 
failure from material, manufacturing, and service induced anomalies 
within the approved life of the part. Applicants must publish a list of 
the life-limited engine parts and the approved life for each part in the 
Airworthiness Limitations Section of the Instructions for Continued 
Airworthiness as required by Sec. 33.4 of this part.
    (b) A manufacturing plan that identifies the specific manufacturing 
constraints necessary to consistently produce each engine life-limited 
part with the attributes required by the engineering plan.
    (c) A service management plan that defines in-service processes for 
maintenance and the limitations to repair for each engine life-limited 
part that will maintain attributes consistent with those required by the 
engineering plan. These processes and limitations will become part of 
the Instructions for Continued Airworthiness.

[Amdt. 33-22, 72 FR 50860, Sept. 4, 2007]



Sec. 33.71  Lubrication system.

    (a) General. Each lubrication system must function properly in the 
flight attitudes and atmospheric conditions in which an aircraft is 
expected to operate.
    (b) Oil strainer or filter. There must be an oil strainer or filter 
through which all of the engine oil flows. In addition:
    (1) Each strainer or filter required by this paragraph that has a 
bypass must be constructed and installed so that oil will flow at the 
normal rate through the rest of the system with the strainer or filter 
element completely blocked.
    (2) The type and degree of filtering necessary for protection of the 
engine oil system against foreign particles in the oil must be 
specified. The applicant must demonstrate that foreign particles passing 
through the specified filtering means do not impair engine oil system 
functioning.
    (3) Each strainer or filter required by this paragraph must have the 
capacity (with respect to operating limitations established for the 
engine) to ensure that engine oil system functioning is not impaired 
with the oil contaminated to a degree (with respect to particle size and 
density) that is greater than that established for the engine in 
paragraph (b)(2) of this section.
    (4) For each strainer or filter required by this paragraph, except 
the strainer or filter at the oil tank outlet, there must be means to 
indicate contamination before it reaches the capacity established in 
accordance with paragraph (b)(3) of this section.
    (5) Any filter bypass must be designed and constructed so that the 
release of collected contaminants is minimized by appropriate location 
of the bypass to ensure that the collected contaminants are not in the 
bypass flow path.
    (6) Each strainer or filter required by this paragraph that has no 
bypass, except the strainer or filter at an oil tank outlet or for a 
scavenge pump,

[[Page 889]]

must have provisions for connection with a warning means to warn the 
pilot of the occurance of contamination of the screen before it reaches 
the capacity established in accordance with paragraph (b)(3) of this 
section.
    (7) Each strainer or filter required by this paragraph must be 
accessible for draining and cleaning.
    (c) Oil tanks. (1) Each oil tank must have an expansion space of not 
less than 10 percent of the tank capacity.
    (2) It must be impossible to inadvertently fill the oil tank 
expansion space.
    (3) Each recessed oil tank filler connection that can retain any 
appreciable quantity of oil must have provision for fitting a drain.
    (4) Each oil tank cap must provide an oil-tight seal. For an 
applicant seeking eligibility for an engine to be installed on an 
airplane approved for ETOPS, the oil tank must be designed to prevent a 
hazardous loss of oil due to an incorrectly installed oil tank cap.
    (5) Each oil tank filler must be marked with the word ``oil.''
    (6) Each oil tank must be vented from the top part of the expansion 
space, with the vent so arranged that condensed water vapor that might 
freeze and obstruct the line cannot accumulate at any point.
    (7) There must be means to prevent entrance into the oil tank or 
into any oil tank outlet, of any object that might obstruct the flow of 
oil through the system.
    (8) There must be a shutoff valve at the outlet of each oil tank, 
unless the external portion of the oil system (including oil tank 
supports) is fireproof.
    (9) Each unpressurized oil tank may not leak when subjected to a 
maximum operating temperature and an internal pressure of 5 p.s.i., and 
each pressurized oil tank must meet the requirements of Sec. 33.64.
    (10) Leaked or spilled oil may not accumulate between the tank and 
the remainder of the engine.
    (11) Each oil tank must have an oil quantity indicator or provisions 
for one.
    (12) If the propeller feathering system depends on engine oil--
    (i) There must be means to trap an amount of oil in the tank if the 
supply becomes depleted due to failure of any part of the lubricating 
system other than the tank itself;
    (ii) The amount of trapped oil must be enough to accomplish the 
feathering opeation and must be available only to the feathering pump; 
and
    (iii) Provision must be made to prevent sludge or other foreign 
matter from affecting the safe operation of the propeller feathering 
system.
    (d) Oil drains. A drain (or drains) must be provided to allow safe 
drainage of the oil system. Each drain must--
    (1) Be accessible; and
    (2) Have manual or automatic means for positive locking in the 
closed position.
    (e) Oil radiators. Each oil radiator must withstand, without 
failure, any vibration, inertia, and oil pressure load to which it is 
subjected during the block tests.

[Amdt. 33-6, 39 FR 35466, Oct. 1, 1974, as amended by Amdt. 33-10, 49 FR 
6852, Feb. 23, 1984; Amdt. 33-21, 72 FR 1877, Jan. 16, 2007; Amdt. 33-
27, 73 FR 55437, Sept. 25, 2008; Amdt. 33-27, 73 FR 57235, Oct. 2, 2008]



Sec. 33.72  Hydraulic actuating systems.

    Each hydraulic actuating system must function properly under all 
conditions in which the engine is expected to operate. Each filter or 
screen must be accessible for servicing and each tank must meet the 
design criteria of Sec. 33.71.

[Amdt. 33-6, 39 FR 35467, Oct. 1, 1974]



Sec. 33.73  Power or thrust response.

    The design and construction of the engine must enable an increase--
    (a) From minimum to rated takeoff power or thrust with the maximum 
bleed air and power extraction to be permitted in an aircraft, without 
overtemperature, surge, stall, or other detrimental factors occurring to 
the engine whenever the power control lever is moved from the minimum to 
the maximum position in not more than 1 second, except that the 
Administrator may allow additional time increments for different regimes 
of control operation requiring control scheduling; and
    (b) From the fixed minimum flight idle power lever position when 
provided, or if not provided, from not more than 15 percent of the rated 
takeoff power or thrust available to 95 percent rated takeoff power or 
thrust in not

[[Page 890]]

over 5 seconds. The 5-second power or thrust response must occur from a 
stabilized static condition using only the bleed air and accessories 
loads necessary to run the engine. This takeoff rating is specified by 
the applicant and need not include thrust augmentation.

[Amdt. 33-1, 36 FR 5493, Mar. 24, 1971]



Sec. 33.74  Continued rotation.

    If any of the engine main rotating systems continue to rotate after 
the engine is shutdown for any reason while in flight, and if means to 
prevent that continued rotation are not provided, then any continued 
rotation during the maximum period of flight, and in the flight 
conditions expected to occur with that engine inoperative, may not 
result in any condition described in Sec. 33.75(g)(2)(i) through (vi) 
of this part.

[Amdt. 33-24, 72 FR 50867, Sept. 4, 2007]



Sec. 33.75  Safety analysis.

    (a) (1) The applicant must analyze the engine, including the control 
system, to assess the likely consequences of all failures that can 
reasonably be expected to occur. This analysis will take into account, 
if applicable:
    (i) Aircraft-level devices and procedures assumed to be associated 
with a typical installation. Such assumptions must be stated in the 
analysis.
    (ii) Consequential secondary failures and latent failures.
    (iii) Multiple failures referred to in paragraph (d) of this section 
or that result in the hazardous engine effects defined in paragraph 
(g)(2) of this section.
    (2) The applicant must summarize those failures that could result in 
major engine effects or hazardous engine effects, as defined in 
paragraph (g) of this section, and estimate the probability of 
occurrence of those effects. Any engine part the failure of which could 
reasonably result in a hazardous engine effect must be clearly 
identified in this summary.
    (3) The applicant must show that hazardous engine effects are 
predicted to occur at a rate not in excess of that defined as extremely 
remote (probability range of 10-7 to 10-9 per 
engine flight hour). Since the estimated probability for individual 
failures may be insufficiently precise to enable the applicant to assess 
the total rate for hazardous engine effects, compliance may be shown by 
demonstrating that the probability of a hazardous engine effect arising 
from an individual failure can be predicted to be not greater than 
10-8 per engine flight hour. In dealing with probabilities of 
this low order of magnitude, absolute proof is not possible, and 
compliance may be shown by reliance on engineering judgment and previous 
experience combined with sound design and test philosophies.
    (4) The applicant must show that major engine effects are predicted 
to occur at a rate not in excess of that defined as remote (probability 
range of 10-5 to 10-7 per engine flight hour).
    (b) The FAA may require that any assumption as to the effects of 
failures and likely combination of failures be verified by test.
    (c) The primary failure of certain single elements cannot be 
sensibly estimated in numerical terms. If the failure of such elements 
is likely to result in hazardous engine effects, then compliance may be 
shown by reliance on the prescribed integrity requirements of Sec. Sec. 
33.15, 33.27, and 33.70 as applicable. These instances must be stated in 
the safety analysis.
    (d) If reliance is placed on a safety system to prevent a failure 
from progressing to hazardous engine effects, the possibility of a 
safety system failure in combination with a basic engine failure must be 
included in the analysis. Such a safety system may include safety 
devices, instrumentation, early warning devices, maintenance checks, and 
other similar equipment or procedures. If items of a safety system are 
outside the control of the engine manufacturer, the assumptions of the 
safety analysis with respect to the reliability of these parts must be 
clearly stated in the analysis and identified in the installation 
instructions under Sec. 33.5 of this part.
    (e) If the safety analysis depends on one or more of the following 
items, those items must be identified in the analysis and appropriately 
substantiated.
    (1) Maintenance actions being carried out at stated intervals. This 
includes the verification of the serviceability of

[[Page 891]]

items that could fail in a latent manner. When necessary to prevent 
hazardous engine effects, these maintenance actions and intervals must 
be published in the instructions for continued airworthiness required 
under Sec. 33.4 of this part. Additionally, if errors in maintenance of 
the engine, including the control system, could lead to hazardous engine 
effects, the appropriate procedures must be included in the relevant 
engine manuals.
    (2) Verification of the satisfactory functioning of safety or other 
devices at pre-flight or other stated periods. The details of this 
satisfactory functioning must be published in the appropriate manual.
    (3) The provisions of specific instrumentation not otherwise 
required.
    (4) Flight crew actions to be specified in the operating 
instructions established under Sec. 33.5.
    (f) If applicable, the safety analysis must also include, but not be 
limited to, investigation of the following:
    (1) Indicating equipment;
    (2) Manual and automatic controls;
    (3) Compressor bleed systems;
    (4) Refrigerant injection systems;
    (5) Gas temperature control systems;
    (6) Engine speed, power, or thrust governors and fuel control 
systems;
    (7) Engine overspeed, overtemperature, or topping limiters;
    (8) Propeller control systems; and
    (9) Engine or propeller thrust reversal systems.
    (g) Unless otherwise approved by the FAA and stated in the safety 
analysis, for compliance with part 33, the following failure definitions 
apply to the engine:
    (1) An engine failure in which the only consequence is partial or 
complete loss of thrust or power (and associated engine services) from 
the engine will be regarded as a minor engine effect.
    (2) The following effects will be regarded as hazardous engine 
effects:
    (i) Non-containment of high-energy debris;
    (ii) Concentration of toxic products in the engine bleed air 
intended for the cabin sufficient to incapacitate crew or passengers;
    (iii) Significant thrust in the opposite direction to that commanded 
by the pilot;
    (iv) Uncontrolled fire;
    (v) Failure of the engine mount system leading to inadvertent engine 
separation;
    (vi) Release of the propeller by the engine, if applicable; and
    (vii) Complete inability to shut the engine down.
    (3) An effect whose severity falls between those effects covered in 
paragraphs (g)(1) and (g)(2) of this section will be regarded as a major 
engine effect.

[Amdt. 33-24, 72 FR 50867, Sept. 4, 2007]



Sec. 33.76  Bird ingestion.

    (a) General. Compliance with paragraphs (b), (c), and (d) of this 
section shall be in accordance with the following:
    (1) Except as specified in paragraph (d) of this section, all 
ingestion tests must be conducted with the engine stabilized at no less 
than 100-percent takeoff power or thrust, for test day ambient 
conditions prior to the ingestion. In addition, the demonstration of 
compliance must account for engine operation at sea level takeoff 
conditions on the hottest day that a minimum engine can achieve maximum 
rated takeoff thrust or power.
    (2) The engine inlet throat area as used in this section to 
determine the bird quantity and weights will be established by the 
applicant and identified as a limitation in the installation 
instructions required under Sec. 33.5.
    (3) The impact to the front of the engine from the large single 
bird, the single largest medium bird which can enter the inlet, and the 
large flocking bird must be evaluated. Applicants must show that the 
associated components when struck under the conditions prescribed in 
paragraphs (b), (c) or (d) of this section, as applicable, will not 
affect the engine to the extent that the engine cannot comply with the 
requirements of paragraphs (b)(3), (c)(6) and (d)(4) of this section.
    (4) For an engine that incorporates an inlet protection device, 
compliance with this section shall be established with the device 
functioning. The engine approval will be endorsed to show

[[Page 892]]

that compliance with the requirements has been established with the 
device functioning.
    (5) Objects that are accepted by the Administrator may be 
substituted for birds when conducting the bird ingestion tests required 
by paragraphs (b), (c) and (d) of this section.
    (6) If compliance with the requirements of this section is not 
established, the engine type certification documentation will show that 
the engine shall be limited to aircraft installations in which it is 
shown that a bird cannot strike the engine, or be ingested into the 
engine, or adversely restrict airflow into the engine.
    (b) Large single bird.Compliance with the large bird ingestion 
requirements shall be in accordance with the following:
    (1) The large bird ingestion test shall be conducted using one bird 
of a weight determined from Table 1 aimed at the most critical exposed 
location on the first stage rotor blades and ingested at a bird speed of 
200-knots for engines to be installed on airplanes, or the maximum 
airspeed for normal rotorcraft flight operations for engines to be 
installed on rotorcraft.
    (2) Power lever movement is not permitted within 15 seconds 
following ingestion of the large bird.
    (3) Ingestion of a single large bird tested under the conditions 
prescribed in this section may not result in any condition described in 
Sec. 33.75(g)(2) of this part.
    (4) Compliance with the large bird ingestion requirements of this 
paragraph may be shown by demonstrating that the requirements of Sec. 
33.94(a) constitute a more severe demonstration of blade containment and 
rotor unbalance than the requirements of this paragraph.

         Table 1 to Sec. 33.76--Large Bird Weight Requirements
------------------------------------------------------------------------
   Engine Inlet Throat Area (A)--Square-
          meters (square-inches)                Bird weight kg. (lb.)
------------------------------------------------------------------------
1.35 (2,092)A..................  1.85 (4.07) minimum, unless
                                             a smaller bird is
                                             determined to be a more
                                             severe demonstration.
1.35 (2,092)<=A<3.90 (6,045)..............  2.75 (6.05)
3.90 (6,045)<=A...........................  3.65 (8.03)
------------------------------------------------------------------------

    (c) Small and medium flocking bird. Compliance with the small and 
medium bird ingestion requirements shall be in accordance with the 
following:
    (1) Analysis or component test, or both, acceptable to the 
Administrator, shall be conducted to determine the critical ingestion 
parameters affecting power loss and damage. Critical ingestion 
parameters shall include, but are not limited to, the effects of bird 
speed, critical target location, and first stage rotor speed. The 
critical bird ingestion speed should reflect the most critical condition 
within the range of airspeeds used for normal flight operations up to 
1,500 feet above ground level, but not less than V1 minimum 
for airplanes.
    (2) Medium bird engine tests shall be conducted so as to simulate a 
flock encounter, and will use the bird weights and quantities specified 
in Table 2. When only one bird is specified, that bird will be aimed at 
the engine core primary flow path; the other critical locations on the 
engine face area must be addressed, as necessary, by appropriate tests 
or analysis, or both. When two or more birds are specified in Table 2, 
the largest of those birds must be aimed at the engine core primary flow 
path, and a second bird must be aimed at the most critical exposed 
location on the first stage rotor blades. Any remaining birds must be 
evenly distributed over the engine face area.
    (3) In addition, except for rotorcraft engines, it must also be 
substantiated by appropriate tests or analysis or both, that when the 
full fan assembly is subjected to the ingestion of the quantity and 
weights of bird from Table 3, aimed at the fan assembly's most critical 
location outboard of the primary core flowpath, and in accordance with 
the applicable test conditions of this paragraph, that the engine can 
comply with the acceptance criteria of this paragraph.
    (4) A small bird ingestion test is not required if the prescribed 
number of medium birds pass into the engine rotor blades during the 
medium bird test.
    (5) Small bird ingestion tests shall be conducted so as to simulate 
a flock encounter using one 85 gram (0.187 lb.) bird for each 0.032 
square-meter (49.6 square-inches) of inlet area, or fraction

[[Page 893]]

thereof, up to a maximum of 16 birds. The birds will be aimed so as to 
account for any critical exposed locations on the first stage rotor 
blades, with any remaining birds evenly distributed over the engine face 
area.
    (6) Ingestion of small and medium birds tested under the conditions 
prescribed in this paragraph may not cause any of the following:
    (i) More than a sustained 25-percent power or thrust loss;
    (ii) The engine to be shut down during the required run-on 
demonstration prescribed in paragraphs (c)(7) or (c)(8) of this section;
    (iii) The conditions defined in paragraph (b)(3) of this section.
    (iv) Unacceptable deterioration of engine handling characteristics.
    (7) Except for rotorcraft engines, the following test schedule shall 
be used:
    (i) Ingestion so as to simulate a flock encounter, with 
approximately 1 second elapsed time from the moment of the first bird 
ingestion to the last.
    (ii) Followed by 2 minutes without power lever movement after the 
ingestion.
    (iii) Followed by 3 minutes at 75-percent of the test condition.
    (iv) Followed by 6 minutes at 60-percent of the test condition.
    (v) Followed by 6 minutes at 40-percent of the test condition.
    (vi) Followed by 1 minute at approach idle.
    (vii) Followed by 2 minutes at 75-percent of the test condition.
    (viii) Followed by stabilizing at idle and engine shut down.
    (ix) The durations specified are times at the defined conditions 
with the power being changed between each condition in less than 10 
seconds.
    (8) For rotorcraft engines, the following test schedule shall be 
used:
    (i) Ingestion so as to simulate a flock encounter within 
approximately 1 second elapsed time between the first ingestion and the 
last.
    (ii) Followed by 3 minutes at 75-percent of the test condition.
    (iii) Followed by 90 seconds at descent flight idle.
    (iv) Followed by 30 seconds at 75-percent of the test condition.
    (v) Followed by stabilizing at idle and engine shut down.
    (vi) The durations specified are times at the defined conditions 
with the power being changed between each condition in less than 10 
seconds.
    (9) Engines intended for use in multi-engine rotorcraft are not 
required to comply with the medium bird ingestion portion of this 
section, providing that the appropriate type certificate documentation 
is so endorsed.
    (10) If any engine operating limit(s) is exceeded during the initial 
2 minutes without power lever movement, as provided by paragraph 
(c)(7)(ii) of this section, then it shall be established that the limit 
exceedence will not result in an unsafe condition.

    Table 2 to Sec. 33.76--Medium Flocking Bird Weight and Quantity
                              Requirements
------------------------------------------------------------------------
 Engine Inlet Throat Area (A)--                         Bird weight kg.
  Square-meters (square-inches)      Bird quantity           (lb.)
------------------------------------------------------------------------
0.05 (77.5)A.........  none..............  ..................
0.05 (77.5)<=A <0.10 (155)......  1.................  0.35 (0.77)
0.10 (155)<=A <0.20 (310).......  1.................  0.45 (0.99)
0.20 (310)<=A <0.40 (620).......  2.................  0.45 (0.99)
0.40 (620)<=A <0.60 (930).......  2.................  0.70 (1.54)
0.60 (930)<=A <1.00 (1,550).....  3.................  0.70 (1.54)
1.00 (1,550)<=A <1.35 (2,092)...  4.................  0.70 (1.54)
1.35 (2,092)<=A <1.70 (2,635)...  1.................  1.15 (2.53)
                                  plus 3............  0.70 (1.54)
1.70 (2,635)<=A <2.10 (3,255)...  1.................  1.15 (2.53)
                                  plus 4............  0.70 (1.54)
2.10 (3,255)<=A <2.50 (3,875)...  1.................  1.15 (2.53)
                                  plus 5............  0.70 (1.54)
2.50 (3,875)<=A <3.90 (6045)....  1.................  1.15 (2.53)
                                  plus 6............  0.70 (1.54)
3.90 (6045)<=A <4.50 (6975).....  3.................  1.15 (2.53)
4.50 (6975)<=A..................  4.................  1.15 (2.53)
------------------------------------------------------------------------


        Table 3 to Sec. 33.76--Additional Integrity Assessment
------------------------------------------------------------------------
 Engine Inlet Throat Area (A)--                         Bird weight kg.
  square-meters (square-inches)      Bird quantity           (lb.)
------------------------------------------------------------------------
1.35 (2,092)A........  none..............  ..................
1.35 (2,092)<=A <2.90 (4,495)...  1.................  1.15 (2.53)
2.90 (4,495)<=A <3.90 (6,045)...  2.................  1.15 (2.53)
3.90 (6,045)<=A.................  1.................  1.15 (2.53)
                                  plus 6............  0.70 (1.54)
------------------------------------------------------------------------

    (d) Large flocking bird. An engine test will be performed as 
follows:
    (1) Large flocking bird engine tests will be performed using the 
bird mass and weights in Table 4, and ingested at a bird speed of 200 
knots.
    (2) Prior to the ingestion, the engine must be stabilized at no less 
than the mechanical rotor speed of the first exposed stage or stages 
that, on a standard day, would produce 90 percent of

[[Page 894]]

the sea level static maximum rated takeoff power or thrust.
    (3) The bird must be targeted on the first exposed rotating stage or 
stages at a blade airfoil height of not less than 50 percent measured at 
the leading edge.
    (4) Ingestion of a large flocking bird under the conditions 
prescribed in this paragraph must not cause any of the following:
    (i) A sustained reduction of power or thrust to less than 50 percent 
of maximum rated takeoff power or thrust during the run-on segment 
specified under paragraph (d)(5)(i) of this section.
    (ii) Engine shutdown during the required run-on demonstration 
specified in paragraph (d)(5) of this section.
    (iii) The conditions specified in paragraph (b)(3) of this section.
    (5) The following test schedule must be used:
    (i) Ingestion followed by 1 minute without power lever movement.
    (ii) Followed by 13 minutes at not less than 50 percent of maximum 
rated takeoff power or thrust.
    (iii) Followed by 2 minutes between 30 and 35 percent of maximum 
rated takeoff power or thrust.
    (iv) Followed by 1 minute with power or thrust increased from that 
set in paragraph (d)(5)(iii) of this section, by between 5 and 10 
percent of maximum rated takeoff power or thrust.
    (v) Followed by 2 minutes with power or thrust reduced from that set 
in paragraph (d)(5)(iv) of this section, by between 5 and 10 percent of 
maximum rated takeoff power or thrust.
    (vi) Followed by a minimum of 1 minute at ground idle then engine 
shutdown. The durations specified are times at the defined conditions. 
Power lever movement between each condition will be 10 seconds or less, 
except that power lever movements allowed within paragraph (d)(5)(ii) of 
this section are not limited, and for setting power under paragraph 
(d)(5)(iii) of this section will be 30 seconds or less.
    (6) Compliance with the large flocking bird ingestion requirements 
of this paragraph (d) may also be demonstrated by:
    (i) Incorporating the requirements of paragraph (d)(4) and (d)(5) of 
this section, into the large single bird test demonstration specified in 
paragraph (b)(1) of this section; or
    (ii) Use of an engine subassembly test at the ingestion conditions 
specified in paragraph (b)(1) of this section if:
    (A) All components critical to complying with the requirements of 
paragraph (d) of this section are included in the subassembly test;
    (B) The components of paragraph (d)(6)(ii)(A) of this section are 
installed in a representative engine for a run-on demonstration in 
accordance with paragraphs (d)(4) and (d)(5) of this section; except 
that section (d)(5)(i) is deleted and section (d)(5)(ii) must be 14 
minutes in duration after the engine is started and stabilized; and
    (C) The dynamic effects that would have been experienced during a 
full engine ingestion test can be shown to be negligible with respect to 
meeting the requirements of paragraphs (d)(4) and (d)(5) of this 
section.
    (7) Applicants must show that an unsafe condition will not result if 
any engine operating limit is exceeded during the run-on period.

      Table 4 to Sec. 33.76--Large Flocking Bird Mass and Weight
------------------------------------------------------------------------
                                                              Bird mass
 Engine inlet throat area (square meters/square     Bird     and weight
                     inches)                      quantity   (kg (lbs))
------------------------------------------------------------------------
A <2.50 (3875)..................................      none
2.50 (3875) <=A <3.50 (5425)....................         1   1.85 (4.08)
3.50 (5425) <=A <3.90 (6045)....................         1   2.10 (4.63)
3.90 (6045) <=A.................................         1   2.50 (5.51)
------------------------------------------------------------------------


[Doc. No. FAA-1998-4815, 65 FR 55854, Sept. 14, 2000, as amended by 
Amdt. 33-20, 68 FR 75391, Dec. 31, 2003; Amdt. 33-24, 72 FR 50868, Sept. 
4, 2007; Amdt. 33-23, 72 FR 58974, Oct. 17, 2007]



Sec. 33.77  Foreign object ingestion--ice.

    (a)-(b) [Reserved]
    (c) Ingestion of ice under the conditions of paragraph (e) of this 
section may not--
    (1) Cause a sustained power or thrust loss; or
    (2) Require the engine to be shutdown.
    (d) For an engine that incorporates a protection device, compliance 
with this section need not be demonstrated with respect to foreign 
objects to be ingested under the conditions prescribed

[[Page 895]]

in paragraph (e) of this section if it is shown that--
    (1) Such foreign objects are of a size that will not pass through 
the protective device;
    (2) The protective device will withstand the impact of the foreign 
objects; and
    (3) The foreign object, or objects, stopped by the protective device 
will not obstruct the flow of induction air into the engine with a 
resultant sustained reduction in power or thrust greater than those 
values required by paragraph (c) of this section.
    (e) Compliance with paragraph (c) of this section must be shown by 
engine test under the following ingestion conditions:
    (1) Ice quantity will be the maximum accumulation on a typical inlet 
cowl and engine face resulting from a 2-minute delay in actuating the 
anti-icing system; or a slab of ice which is comparable in weight or 
thickness for that size engine.
    (2) The ingestion velocity will simulate ice being sucked into the 
engine inlet.
    (3) Engine operation will be maximum cruise power or thrust.
    (4) The ingestion will simulate a continuous maximum icing encounter 
at 25 degrees Fahrenheit.

[Doc. No. 16919, 49 FR 6852, Feb. 23, 1984, as amended by Amdt. 33-19, 
63 FR 14798, Mar. 26, 1998; 63 FR 53278, Oct. 5, 1998; Amdt. 33-20, 65 
FR 55856, Sept. 14, 2000]

    Effective Date Note: At 79 FR 65537, Nov. 4, 2014, Sec. 33.77 was 
amended by adding paragraph (a) and revising paragraphs (c) introductory 
text, (c)(1), (d), and (e), effective Jan. 5, 2015. For the convenience 
of the user, the added and revised text is set forth as follows:



Sec. 33.77  Foreign object ingestion ice.

    (a) Compliance with the requirements of this section must be 
demonstrated by engine ice ingestion test or by validated analysis 
showing equivalence of other means for demonstrating soft body damage 
tolerance.

                                * * * * *

    (c) Ingestion of ice under the conditions of this section may not--
    (1) Cause an immediate or ultimate unacceptable sustained power or 
thrust loss; or

                                * * * * *

    (d) For an engine that incorporates a protection device, compliance 
with this section need not be demonstrated with respect to ice formed 
forward of the protection device if it is shown that--
    (1) Such ice is of a size that will not pass through the protective 
device;
    (2) The protective device will withstand the impact of the ice; and
    (3) The ice stopped by the protective device will not obstruct the 
flow of induction air into the engine with a resultant sustained 
reduction in power or thrust greater than those values defined by 
paragraph (c) of this section.
    (e) Compliance with the requirements of this section must be 
demonstrated by engine ice ingestion test under the following ingestion 
conditions or by validated analysis showing equivalence of other means 
for demonstrating soft body damage tolerance.
    (1) The minimum ice quantity and dimensions will be established by 
the engine size as defined in Table 1 of this section.
    (2) The ingested ice dimensions are determined by linear 
interpolation between table values, and are based on the actual engine's 
inlet hilite area.
    (3) The ingestion velocity will simulate ice from the inlet being 
sucked into the engine.
    (4) Engine operation will be at the maximum cruise power or thrust 
unless lower power is more critical.

                         Table 1--Minimum Ice Slab Dimensions Based on Engine Inlet Size
----------------------------------------------------------------------------------------------------------------
                                                                     Thickness
              Engine Inlet Hilite area  (sq. inch)                    (inch)       Width  (inch)  Length  (inch)
----------------------------------------------------------------------------------------------------------------
0...............................................................            0.25               0             3.6
80..............................................................            0.25               6             3.6
300.............................................................            0.25              12             3.6
700.............................................................            0.25              12             4.8
2800............................................................            0.35              12             8.5
5000............................................................            0.43              12            11.0
7000............................................................            0.50              12            12.7
7900............................................................            0.50              12            13.4
9500............................................................            0.50              12            14.6
11300...........................................................            0.50              12            15.9
13300...........................................................            0.50              12            17.1
16500...........................................................            0.5               12            18.9

[[Page 896]]

 
20000...........................................................            0.5               12            20.0
----------------------------------------------------------------------------------------------------------------



Sec. 33.78  Rain and hail ingestion.

    (a) All engines. (1) The ingestion of large hailstones (0.8 to 0.9 
specific gravity) at the maximum true air speed, up to 15,000 feet 
(4,500 meters), associated with a representative aircraft operating in 
rough air, with the engine at maximum continuous power, may not cause 
unacceptable mechanical damage or unacceptable power or thrust loss 
after the ingestion, or require the engine to be shut down. One-half the 
number of hailstones shall be aimed randomly over the inlet face area 
and the other half aimed at the critical inlet face area. The hailstones 
shall be ingested in a rapid sequence to simulate a hailstone encounter 
and the number and size of the hailstones shall be determined as 
follows:
    (i) One 1-inch (25 millimeters) diameter hailstone for engines with 
inlet areas of not more than 100 square inches (0.0645 square meters).
    (ii) One 1-inch (25 millimeters) diameter and one 2-inch (50 
millimeters) diameter hailstone for each 150 square inches (0.0968 
square meters) of inlet area, or fraction thereof, for engines with 
inlet areas of more than 100 square inches (0.0645 square meters).
    (2) In addition to complying with paragraph (a)(1) of this section 
and except as provided in paragraph (b) of this section, it must be 
shown that each engine is capable of acceptable operation throughout its 
specified operating envelope when subjected to sudden encounters with 
the certification standard concentrations of rain and hail, as defined 
in appendix B to this part. Acceptable engine operation precludes 
flameout, run down, continued or non-recoverable surge or stall, or loss 
of acceleration and deceleration capability, during any three minute 
continuous period in rain and during any 30 second continuous period in 
hail. It must also be shown after the ingestion that there is no 
unacceptable mechanical damage, unacceptable power or thrust loss, or 
other adverse engine anomalies.
    (b) Engines for rotorcraft. As an alternative to the requirements 
specified in paragraph (a)(2) of this section, for rotorcraft turbine 
engines only, it must be shown that each engine is capable of acceptable 
operation during and after the ingestion of rain with an overall ratio 
of water droplet flow to airflow, by weight, with a uniform distribution 
at the inlet plane, of at least four percent. Acceptable engine 
operation precludes flameout, run down, continued or non-recoverable 
surge or stall, or loss of acceleration and deceleration capability. It 
must also be shown after the ingestion that there is no unacceptable 
mechanical damage, unacceptable power loss, or other adverse engine 
anomalies. The rain ingestion must occur under the following static 
ground level conditions:
    (1) A normal stabilization period at take-off power without rain 
ingestion, followed immediately by the suddenly commencing ingestion of 
rain for three minutes at takeoff power, then
    (2) Continuation of the rain ingestion during subsequent rapid 
deceleration to minimum idle, then
    (3) Continuation of the rain ingestion during three minutes at 
minimum idle power to be certified for flight operation, then
    (4) Continuation of the rain ingestion during subsequent rapid 
acceleration to takeoff power.
    (c) Engines for supersonic airplanes. In addition to complying with 
paragraphs (a)(1) and (a)(2) of this section, a separate test for 
supersonic airplane engines only, shall be conducted with three 
hailstones ingested at supersonic cruise velocity. These hailstones 
shall be aimed at the engine's critical face area, and their ingestion 
must not cause unacceptable mechanical damage or unacceptable power or 
thrust loss after the ingestion or require the engine to be shut down. 
The size of these hailstones shall be determined from

[[Page 897]]

the linear variation in diameter from 1-inch (25 millimeters) at 35,000 
feet (10,500 meters) to \1/4\-inch (6 millimeters) at 60,000 feet 
(18,000 meters) using the diameter corresponding to the lowest expected 
supersonic cruise altitude. Alternatively, three larger hailstones may 
be ingested at subsonic velocities such that the kinetic energy of these 
larger hailstones is equivalent to the applicable supersonic ingestion 
conditions.
    (d) For an engine that incorporates or requires the use of a 
protection device, demonstration of the rain and hail ingestion 
capabilities of the engine, as required in paragraphs (a), (b), and (c) 
of this section, may be waived wholly or in part by the Administrator if 
the applicant shows that:
    (1) The subject rain and hail constituents are of a size that will 
not pass through the protection device;
    (2) The protection device will withstand the impact of the subject 
rain and hail constituents; and
    (3) The subject of rain and hail constituents, stopped by the 
protection device, will not obstruct the flow of induction air into the 
engine, resulting in damage, power or thrust loss, or other adverse 
engine anomalies in excess of what would be accepted in paragraphs (a), 
(b), and (c) of this section.

[Doc. No. 28652, 63 FR 14799, Mar. 26, 1998]



Sec. 33.79  Fuel burning thrust augmentor.

    Each fuel burning thrust augmentor, including the nozzle, must--
    (a) Provide cutoff of the fuel burning thrust augmentor;
    (b) Permit on-off cycling;
    (c) Be controllable within the intended range of operation;
    (d) Upon a failure or malfunction of augmentor combustion, not cause 
the engine to lose thrust other than that provided by the augmentor; and
    (e) Have controls that function compatibly with the other engine 
controls and automatically shut off augmentor fuel flow if the engine 
rotor speed drops below the minimum rotational speed at which the 
augmentor is intended to function.

[Amdt. 33-6, 39 FR 35468, Oct. 1, 1974]



             Subpart F_Block Tests; Turbine Aircraft Engines



Sec. 33.81  Applicability.

    This subpart prescribes the block tests and inspections for turbine 
engines.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-6, 39 
FR 35468, Oct. 1, 1974]



Sec. 33.82  General.

    Before each endurance test required by this subpart, the adjustment 
setting and functioning characteristic of each component having an 
adjustment setting and a functioning characteristic that can be 
established independent of installation on the engine must be 
established and recorded.

[Amdt. 36-6, 39 FR 35468, Oct. 1, 1974]



Sec. 33.83  Vibration test.

    (a) Each engine must undergo vibration surveys to establish that the 
vibration characteristics of those components that may be subject to 
mechanically or aerodynamically induced vibratory excitations are 
acceptable throughout the declared flight envelope. The engine surveys 
shall be based upon an appropriate combination of experience, analysis, 
and component test and shall address, as a minimum, blades, vanes, rotor 
discs, spacers, and rotor shafts.
    (b) The surveys shall cover the ranges of power or thrust, and both 
the physical and corrected rotational speeds for each rotor system, 
corresponding to operations throughout the range of ambient conditions 
in the declared flight envelope, from the minimum rotational speed up to 
103 percent of the maximum physical and corrected rotational speed 
permitted for rating periods of two minutes or longer, and up to 100 
percent of all other permitted physical and corrected rotational speeds, 
including those that are overspeeds. If there is any indication of a 
stress peak arising at the highest of those required physical or 
corrected rotational speeds, the surveys shall be extended sufficiently 
to reveal the maximum stress values present, except that the extension 
need

[[Page 898]]

not cover more than a further 2 percentage points increase beyond those 
speeds.
    (c) Evaluations shall be made of the following:
    (1) The effects on vibration characteristics of operating with 
scheduled changes (including tolerances) to variable vane angles, 
compressor bleeds, accessory loading, the most adverse inlet air flow 
distortion pattern declared by the manufacturer, and the most adverse 
conditions in the exhaust duct(s); and
    (2) The aerodynamic and aeromechanical factors which might induce or 
influence flutter in those systems susceptible to that form of 
vibration.
    (d) Except as provided by paragraph (e) of this section, the 
vibration stresses associated with the vibration characteristics 
determined under this section, when combined with the appropriate steady 
stresses, must be less than the endurance limits of the materials 
concerned, after making due allowances for operating conditions for the 
permitted variations in properties of the materials. The suitability of 
these stress margins must be justified for each part evaluated. If it is 
determined that certain operating conditions, or ranges, need to be 
limited, operating and installation limitations shall be established.
    (e) The effects on vibration characteristics of excitation forces 
caused by fault conditions (such as, but not limited to, out-of balance, 
local blockage or enlargement of stator vane passages, fuel nozzle 
blockage, incorrectly schedule compressor variables, etc.) shall be 
evaluated by test or analysis, or by reference to previous experience 
and shall be shown not to create a hazardous condition.
    (f) Compliance with this section shall be substantiated for each 
specific installation configuration that can affect the vibration 
characteristics of the engine. If these vibration effects cannot be 
fully investigated during engine certification, the methods by which 
they can be evaluated and methods by which compliance can be shown shall 
be substantiated and defined in the installation instructions required 
by Sec. 33.5.

[Doc. No. 28107, 61 FR 28433, June 4, 1996, as amended by Amdt. 33-33, 
77 FR 39624, July 5, 2012; 77 FR 58301, Sept. 20, 2012]



Sec. 33.84  Engine overtorque test.

    (a) If approval of a maximum engine overtorque is sought for an 
engine incorporating a free power turbine, compliance with this section 
must be demonstrated by testing.
    (1) The test may be run as part of the endurance test requirement of 
Sec. 33.87. Alternatively, tests may be performed on a complete engine 
or equivalent testing on individual groups of components.
    (2) Upon conclusion of tests conducted to show compliance with this 
section, each engine part or individual groups of components must meet 
the requirements of Sec. 33.93(a)(1) and (a)(2).
    (b) The test conditions must be as follows:
    (1) A total of 15 minutes run at the maximum engine overtorque to be 
approved. This may be done in separate runs, each being of at least 2\1/
2\ minutes duration.
    (2) A power turbine rotational speed equal to the highest speed at 
which the maximum overtorque can occur in service. The test speed may 
not be more than the limit speed of take-off or OEI ratings longer than 
2 minutes.
    (3) For engines incorporating a reduction gearbox, a gearbox oil 
temperature equal to the maximum temperature when the maximum engine 
overtorque could occur in service; and for all other engines, an oil 
temperature within the normal operating range.
    (4) A turbine entry gas temperature equal to the maximum steady 
state temperature approved for use during periods longer than 20 seconds 
when operating at conditions not associated with 30-second or 2 minutes 
OEI ratings. The requirement to run the test at the maximum approved 
steady state temperature may be waived by the FAA if the applicant can 
demonstrate that other testing provides substantiation of the 
temperature effects when considered in combination with the

[[Page 899]]

other parameters identified in paragraphs (b)(1), (b)(2) and (b)(3) of 
this section.

[Doc. No. 2007-28502, 74 FR 45310, Sept. 2, 2009]



Sec. 33.85  Calibration tests.

    (a) Each engine must be subjected to those calibration tests 
necessary to establish its power characteristics and the conditions for 
the endurance test specified Sec. 33.87. The results of the power 
characteristics calibration tests form the basis for establishing the 
characteristics of the engine over its entire operating range of speeds, 
pressures, temperatures, and altitudes. Power ratings are based upon 
standard atmospheric conditions with no airbleed for aircraft services 
and with only those accessories installed which are essential for engine 
functioning.
    (b) A power check at sea level conditions must be accomplished on 
the endurance test engine after the endurance test and any change in 
power characteristics which occurs during the endurance test must be 
determined. Measurements taken during the final portion of the endurance 
test may be used in showing compliance with the requirements of this 
paragraph.
    (c) In showing compliance with this section, each condition must 
stabilize before measurements are taken, except as permitted by 
paragraph (d) of this section.
    (d) In the case of engines having 30-second OEI, and 2-minute OEI 
ratings, measurements taken during the applicable endurance test 
prescribed in Sec. 33.87(f) (1) through (8) may be used in showing 
compliance with the requirements of this section for these OEI ratings.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-6, 39 
FR 35468, Oct. 1, 1974; Amdt. 33-18, 61 FR 31328, June 19, 1996]



Sec. 33.87  Endurance test.

    (a) General. Each engine must be subjected to an endurance test that 
includes a total of at least 150 hours of operation and, depending upon 
the type and contemplated use of the engine, consists of one of the 
series of runs specified in paragraphs (b) through (g) of this section, 
as applicable. For engines tested under paragraphs (b), (c), (d), (e) or 
(g) of this section, the prescribed 6-hour test sequence must be 
conducted 25 times to complete the required 150 hours of operation. 
Engines for which the 30-second OEI and 2-minute OEI ratings are desired 
must be further tested under paragraph (f) of this section. The 
following test requirements apply:
    (1) The runs must be made in the order found appropriate by the FAA 
for the particular engine being tested.
    (2) Any automatic engine control that is part of the engine must 
control the engine during the endurance test except for operations where 
automatic control is normally overridden by manual control or where 
manual control is otherwise specified for a particular test run.
    (3) Except as provided in paragraph (a)(5) of this section, power or 
thrust, gas temperature, rotor shaft rotational speed, and, if limited, 
temperature of external surfaces of the engine must be at least 100 
percent of the value associated with the particular engine operation 
being tested. More than one test may be run if all parameters cannot be 
held at the 100 percent level simultaneously.
    (4) The runs must be made using fuel, lubricants and hydraulic fluid 
which conform to the specifications specified in complying with Sec. 
33.7(c).
    (5) Maximum air bleed for engine and aircraft services must be used 
during at least one-fifth of the runs, except for the test required 
under paragraph (f) of this section, provided the validity of the test 
is not compromised. However, for these runs, the power or thrust or the 
rotor shaft rotational speed may be less than 100 percent of the value 
associated with the particular operation being tested if the FAA finds 
that the validity of the endurance test is not compromised.
    (6) Each accessory drive and mounting attachment must be loaded in 
accordance with paragraphs (a)(6)(i) and (ii) of this section, except as 
permitted by paragraph (a)(6)(iii) of this section for the test required 
under paragraph (f) of this section.
    (i) The load imposed by each accessory used only for aircraft 
service must

[[Page 900]]

be the limit load specified by the applicant for the engine drive and 
attachment point during rated maximum continuous power or thrust and 
higher output.
    (ii) The endurance test of any accessory drive and mounting 
attachment under load may be accomplished on a separate rig if the 
validity of the test is confirmed by an approved analysis.
    (iii) The applicant is not required to load the accessory drives and 
mounting attachments when running the tests under paragraphs (f)(1) 
through (f)(8) of this section if the applicant can substantiate that 
there is no significant effect on the durability of any accessory drive 
or engine component. However, the applicant must add the equivalent 
engine output power extraction from the power turbine rotor assembly to 
the engine shaft output.
    (7) During the runs at any rated power or thrust the gas temperature 
and the oil inlet temperature must be maintained at the limiting 
temperature except where the test periods are not longer than 5 minutes 
and do not allow stabilization. At least one run must be made with fuel, 
oil, and hydraulic fluid at the minimum pressure limit and at least one 
run must be made with fuel, oil, and hydraulic fluid at the maximum 
pressure limit with fluid temperature reduced as necessary to allow 
maximum pressure to be attained.
    (8) If the number of occurrences of either transient rotor shaft 
overspeed, transient gas overtemperature or transient engine overtorque 
is limited, that number of the accelerations required by paragraphs (b) 
through (g) of this section must be made at the limiting overspeed, 
overtemperature or overtorque. If the number of occurrences is not 
limited, half the required accelerations must be made at the limiting 
overspeed, overtemperature or overtorque.
    (9) For each engine type certificated for use on supersonic aircraft 
the following additional test requirements apply:
    (i) To change the thrust setting, the power control lever must be 
moved from the initial position to the final position in not more than 
one second except for movements into the fuel burning thrust augmentor 
augmentation position if additional time to confirm ignition is 
necessary.
    (ii) During the runs at any rated augmented thrust the hydraulic 
fluid temperature must be maintained at the limiting temperature except 
where the test periods are not long enough to allow stabilization.
    (iii) During the simulated supersonic runs the fuel temperature and 
induction air temperature may not be less than the limiting temperature.
    (iv) The endurance test must be conducted with the fuel burning 
thrust augmentor installed, with the primary and secondary exhaust 
nozzles installed, and with the variable area exhaust nozzles operated 
during each run according to the methods specified in complying with 
Sec. 33.5(b).
    (v) During the runs at thrust settings for maximum continuous thrust 
and percentages thereof, the engine must be operated with the inlet air 
distortion at the limit for those thrust settings.
    (b) Engines other than certain rotorcraft engines. For each engine 
except a rotorcraft engine for which a rating is desired under paragraph 
(c), (d), or (e) of this section, the applicant must conduct the 
following runs:
    (1) Takeoff and idling. One hour of alternate five-minute periods at 
rated takeoff power or thrust and at idling power or thrust. The 
developed powers or thrusts at takeoff and idling conditions and their 
corresponding rotor speed and gas temperature conditions must be as 
established by the power control in accordance with the schedule 
established by the applicant. The applicant may, during any one period, 
manually control the rotor speed, power, or thrust while taking data to 
check performance. For engines with augmented takeoff power ratings that 
involve increases in turbine inlet temperature, rotor speed, or shaft 
power, this period of running at takeoff must be at the augmented 
rating. For engines with augmented takeoff power ratings that do not 
materially increase operating severity, the amount of running conducted 
at the augmented rating is determined by the FAA. In changing the power 
setting after each period, the power-control lever must be

[[Page 901]]

moved in the manner prescribed in paragraph (b)(5) of this section.
    (2) Rated maximum continuous and takeoff power or thrust. Thirty 
minutes at--
    (i) Rated maximum continuous power or thrust during fifteen of the 
twenty-five 6-hour endurance test cycles; and
    (ii) Rated takeoff power or thrust during ten of the twenty-five 6-
hour endurance test cycles.
    (3) Rated maximum continuous power or thrust. One hour and 30 
minutes at rated maximum continuous power or thrust.
    (4) Incremental cruise power or thrust. Two hours and 30 minutes at 
the successive power lever positions corresponding to at least 15 
approximately equal speed and time increments between maximum continuous 
engine rotational speed and ground or minimum idle rotational speed. For 
engines operating at constant speed, the thrust and power may be varied 
in place of speed. If there is significant peak vibration anywhere 
between ground idle and maximum continuous conditions, the number of 
increments chosen may be changed to increase the amount of running made 
while subject to the peak vibrations up to not more than 50 percent of 
the total time spent in incremental running.
    (5) Acceleration and deceleration runs. 30 minutes of accelerations 
and decelerations, consisting of six cycles from idling power or thrust 
to rated takeoff power or thrust and maintained at the takeoff power 
lever position for 30 seconds and at the idling power lever position for 
approximately four and one-half minutes. In complying with this 
paragraph, the power-control lever must be moved from one extreme 
position to the other in not more than one second, except that, if 
different regimes of control operations are incorporated necessitating 
scheduling of the power-control lever motion in going from one extreme 
position to the other, a longer period of time is acceptable, but not 
more than two seconds.
    (6) Starts. One hundred starts must be made, of which 25 starts must 
be preceded by at least a two-hour engine shutdown. There must be at 
least 10 false engine starts, pausing for the applicant's specified 
minimum fuel drainage time, before attempting a normal start. There must 
be at least 10 normal restarts with not longer than 15 minutes since 
engine shutdown. The remaining starts may be made after completing the 
150 hours of endurance testing.
    (c) Rotorcraft engines for which a 30-minute OEI power rating is 
desired. For each rotorcraft engine for which a 30-minute OEI power 
rating is desired, the applicant must conduct the following series of 
tests:
    (1) Takeoff and idling. One hour of alternate 5-minute periods at 
rated takeoff power and at idling power. The developed powers at takeoff 
and idling conditions and their corresponding rotor speed and gas 
temperature conditions must be as established by the power control in 
accordance with the schedule established by the applicant. During any 
one period, the rotor speed and power may be controlled manually while 
taking data to check performance. For engines with augmented takeoff 
power ratings that involve increases in turbine inlet temperature, rotor 
speed, or shaft power, this period of running at rated takeoff power 
must be at the augmented power rating. In changing the power setting 
after each period, the power control lever must be moved in the manner 
prescribed in paragraph (c)(6) of this section.
    (2) Rated maximum continuous and takeoff power. Thirty minutes at--
    (i) Rated maximum continuous power during fifteen of the twenty-five 
6-hour endurance test cycles; and
    (ii) Rated takeoff power during ten of the twenty-five 6-hour 
endurance test cycles.
    (3) Rated maximum continuous power. One hour at rated maximum 
continuous power.
    (4) Rated 30-minute OEI power. Thirty minutes at rated 30-minute OEI 
power.
    (5) Incremental cruise power. Two hours and 30 minutes at the 
successive power lever positions corresponding with not less than 15 
approximately equal speed and time increments between maximum continuous 
engine rotational speed and ground or minimum idle rotational speed. For 
engines operating at constant speed, power may be varied in place of 
speed. If there are significant peak vibrations anywhere

[[Page 902]]

between ground idle and maximum continuous conditions, the number of 
increments chosen must be changed to increase the amount of running 
conducted while subject to peak vibrations up to not more than 50 
percent of the total time spent in incremental running.
    (6) Acceleration and deceleration runs. Thirty minutes of 
accelerations and decelerations, consisting of six cycles from idling 
power to rated takeoff power and maintained at the takeoff power lever 
position for 30 seconds and at the idling power lever position for 
approximately 4\1/2\ minutes. In complying with this paragraph, the 
power control lever must be moved from one extreme position to the other 
in not more than one second. If, however, different regimes of control 
operations are incorporated that necessitate scheduling of the power 
control lever motion from one extreme position to the other, then a 
longer period of time is acceptable, but not more than two seconds.
    (7) Starts. One hundred starts, of which 25 starts must be preceded 
by at least a two-hour engine shutdown. There must be at least 10 false 
engine starts, pausing for the applicant's specified minimum fuel 
drainage time, before attempting a normal start. There must be at least 
10 normal restarts not more than 15 minutes after engine shutdown. The 
remaining starts may be made after completing the 150 hours of endurance 
testing.
    (d) Rotorcraft engines for which a continuous OEI rating is desired. 
For each rotorcraft engine for which a continuous OEI power rating is 
desired, the applicant must conduct the following series of tests:
    (1) Takeoff and idling. One hour of alternate 5-minute periods at 
rated takeoff power and at idling power. The developed powers at takeoff 
and idling conditions and their corresponding rotor speed and gas 
temperature conditions must be as established by the power control in 
accordance with the schedule established by the applicant. During any 
one period the rotor speed and power may be controlled manually while 
taking data to check performance. For engines with augmented takeoff 
power ratings that involve increases in turbine inlet temperature, rotor 
speed, or shaft power, this period of running at rated takeoff power 
must be at the augmented power rating. In changing the power setting 
after each period, the power control lever must be moved in the manner 
prescribed in paragraph (d)(6) of this section.
    (2) Rated maximum continuous and takeoff power. Thirty minutes at--
    (i) Rated maximum continuous power during fifteen of the twenty-five 
6-hour endurance test cycles; and
    (ii) Rated takeoff power during ten of the twenty-five 6-hour 
endurance test cycles.
    (3) Rated continuous OEI power. One hour at rated continuous OEI 
power.
    (4) Rated maximum continuous power. One hour at rated maximum 
continuous power.
    (5) Incremental cruise power. Two hours at the successive power 
lever positions corresponding with not less than 12 approximately equal 
speed and time increments between maximum continuous engine rotational 
speed and ground or minimum idle rotational speed. For engines operating 
at constant speed, power may be varied in place of speed. If there are 
significant peak vibrations anywhere between ground idle and maximum 
continuous conditions, the number of increments chosen must be changed 
to increase the amount of running conducted while being subjected to the 
peak vibrations up to not more than 50 percent of the total time spent 
in incremental running.
    (6) Acceleration and deceleration runs. Thirty minutes of 
accelerations and decelerations, consisting of six cycles from idling 
power to rated takeoff power and maintained at the takeoff power lever 
position for 30 seconds and at the idling power lever position for 
approximately 4\1/2\ minutes. In complying with this paragraph, the 
power control lever must be moved from one extreme position to the other 
in not more than 1 second, except that if different regimes of control 
operations are incorporated necessitating scheduling of the power 
control lever motion in going from one extreme position to the other, a 
longer period of time is acceptable, but not more than 2 seconds.

[[Page 903]]

    (7) Starts. One hundred starts, of which 25 starts must be preceded 
by at least a 2-hour engine shutdown. There must be at least 10 false 
engine starts, pausing for the applicant's specified minimum fuel 
drainage time, before attempting a normal start. There must be at least 
10 normal restarts with not longer than 15 minutes since engine 
shutdown. The remaining starts may be made after completing the 150 
hours of endurance testing.
    (e) Rotorcraft engines for which a 2\1/2\-minute OEI power rating is 
desired. For each rotorcraft engine for which a 2\1/2\-minute OEI power 
rating is desired, the applicant must conduct the following series of 
tests:
    (1) Takeoff, 2\1/2\-minute OEI, and idling. One hour of alternate 5-
minute periods at rated takeoff power and at idling power except that, 
during the third and sixth takeoff power periods, only 2\1/2\ minutes 
need be conducted at rated takeoff power, and the remaining 2\1/2\ 
minutes must be conducted at rated 2\1/2\-minute OEI power. The 
developed powers at takeoff, 2\1/2\-minute OEI, and idling conditions 
and their corresponding rotor speed and gas temperature conditions must 
be as established by the power control in accordance with the schedule 
established by the applicant. The applicant may, during any one period, 
control manually the rotor speed and power while taking data to check 
performance. For engines with augmented takeoff power ratings that 
involve increases in turbine inlet temperature, rotor speed, or shaft 
power, this period of running at rated takeoff power must be at the 
augmented rating. In changing the power setting after or during each 
period, the power control lever must be moved in the manner prescribed 
in paragraph (b)(5), (c)(6), or (d)(6) of this section, as applicable.
    (2) The tests required in paragraphs (b)(2) through (b)(6), or 
(c)(2) through (c)(7), or (d)(2) through (d)(7) of this section, as 
applicable, except that in one of the 6-hour test sequences, the last 5 
minutes of the 30 minutes at takeoff power test period of paragraph 
(b)(2) of this section, or of the 30 minutes at 30-minute OEI power test 
period of paragraph (c)(4) of this section, or of the l hour at 
continuous OEI power test period of paragraph (d)(3) of this section, 
must be run at 2\1/2\-minute OEI power.
    (f) Rotorcraft Engines for which 30-second OEI and 2-minute OEI 
ratings are desired. For each rotorcraft engine for which 30-second OEI 
and 2-minute OEI power ratings are desired, and following completion of 
the tests under paragraphs (b), (c), (d), or (e) of this section, the 
applicant may disassemble the tested engine to the extent necessary to 
show compliance with the requirements of Sec. 33.93(a). The tested 
engine must then be reassembled using the same parts used during the 
test runs of paragraphs (b), (c), (d), or (e) of this section, except 
those parts described as consumables in the Instructions for Continued 
Airworthiness. Additionally, the tests required in paragraphs (f)(1) 
through (f)(8) of this section must be run continuously. If a stop 
occurs during these tests, the interrupted sequence must be repeated 
unless the applicant shows that the severity of the test would not be 
reduced if it were continued. The applicant must conduct the following 
test sequence four times, for a total time of not less than 120 minutes:
    (1) Takeoff power. Three minutes at rated takeoff power.
    (2)30-second OEI power. Thirty seconds at rated 30-second OEI power.
    (3) 2-minute OEI power. Two minutes at rated 2-minute OEI power.
    (4) 30-minute OEI power, continuous OEI power, or maximum continuous 
power. Five minutes at whichever is the greatest of rated 30-minute OEI 
power, rated continuous OEI power, or rated maximum continuous power, 
except that, during the first test sequence, this period shall be 65 
minutes. However, where the greatest rated power is 30-minute OEI power, 
that sixty-five minute period shall consist of 30 minutes at 30-minute 
OEI power followed by 35 minutes at whichever is the greater of 
continuous OEI power or maximum continuous power.
    (5) 50 percent takeoff power. One minute at 50 percent takeoff 
power.
    (6) 30-second OEI power. Thirty seconds at rated 30-second OEI 
power.
    (7) 2-minute OEI power. Two minutes at rated 2-minute OEI power.
    (8) Idle. One minute at flight idle.

[[Page 904]]

    (g) Supersonic aircraft engines. For each engine type certificated 
for use on supersonic aircraft the applicant must conduct the following:
    (1) Subsonic test under sea level ambient atmospheric conditions. 
Thirty runs of one hour each must be made, consisting of--
    (i) Two periods of 5 minutes at rated takeoff augmented thrust each 
followed by 5 minutes at idle thrust;
    (ii) One period of 5 minutes at rated takeoff thrust followed by 5 
minutes at not more than 15 percent of rated takeoff thrust;
    (iii) One period of 10 minutes at rated takeoff augmented thrust 
followed by 2 minutes at idle thrust, except that if rated maximum 
continuous augmented thrust is lower than rated takeoff augmented 
thrust, 5 of the 10-minute periods must be at rated maximum continuous 
augmented thrust; and
    (iv) Six periods of 1 minute at rated takeoff augmented thrust each 
followed by 2 minutes, including acceleration and deceleration time, at 
idle thrust.
    (2) Simulated supersonic test. Each run of the simulated supersonic 
test must be preceded by changing the inlet air temperature and pressure 
from that attained at subsonic condition to the temperature and pressure 
attained at supersonic velocity, and must be followed by a return to the 
temperature attained at subsonic condition. Thirty runs of 4 hours each 
must be made, consisting of--
    (i) One period of 30 minutes at the thrust obtained with the power 
control lever set at the position for rated maximum continuous augmented 
thrust followed by 10 minutes at the thrust obtained with the power 
control lever set at the position for 90 percent of rated maximum 
continuous augmented thrust. The end of this period in the first five 
runs must be made with the induction air temperature at the limiting 
condition of transient overtemperature, but need not be repeated during 
the periods specified in paragraphs (g)(2)(ii) through (iv) of this 
section;
    (ii) One period repeating the run specified in paragraph (g)(2)(i) 
of this section, except that it must be followed by 10 minutes at the 
thrust obtained with the power control lever set at the position for 80 
percent of rated maximum continuous augmented thrust;
    (iii) One period repeating the run specified in paragraph (g)(2)(i) 
of this section, except that it must be followed by 10 minutes at the 
thrust obtained with the power control lever set at the position for 60 
percent of rated maximum continuous augmented thrust and then 10 minutes 
at not more than 15 percent of rated takeoff thrust;
    (iv) One period repeating the runs specified in paragraphs (g)(2)(i) 
and (ii) of this section; and
    (v) One period of 30 minutes with 25 of the runs made at the thrust 
obtained with the power control lever set at the position for rated 
maximum continuous augmented thrust, each followed by idle thrust and 
with the remaining 5 runs at the thrust obtained with the power control 
lever set at the position for rated maximum continuous augmented thrust 
for 25 minutes each, followed by subsonic operation at not more than 15 
percent or rated takeoff thrust and accelerated to rated takeoff thrust 
for 5 minutes using hot fuel.
    (3) Starts. One hundred starts must be made, of which 25 starts must 
be preceded by an engine shutdown of at least 2 hours. There must be at 
least 10 false engine starts, pausing for the applicant's specified 
minimum fuel drainage time before attempting a normal start. At least 10 
starts must be normal restarts, each made no later than 15 minutes after 
engine shutdown. The starts may be made at any time, including the 
period of endurance testing.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-3, 32 
FR 3737, Mar. 4, 1967; Amdt. 33-6, 39 FR 35468, Oct. 1, 1974; Amdt. 33-
10, 49 FR 6853, Feb. 23, 1984; Amdt. 33-12, 53 FR 34220, Sept. 2, 1988; 
Amdt. 33-18, 61 FR 31328, June 19, 1996; Amdt. 33-25, 73 FR 48123, Aug. 
18, 2008; Amdt. 33-30, 74 FR 45311, Sept. 2, 2009; Amdt. 33-32, 77 FR 
22187, Apr. 13, 2012]



Sec. 33.88  Engine overtemperature test.

    (a) Each engine must run for 5 minutes at maximum permissible rpm 
with the gas temperature at least 75 [deg]F (42 [deg]C) higher than the 
maximum rating's steady-state operating limit, excluding maximum values 
of rpm and gas temperature associated with the 30-second

[[Page 905]]

OEI and 2-minute OEI ratings. Following this run, the turbine assembly 
must be within serviceable limits.
    (b) In addition to the test requirements in paragraph (a) of this 
section, each engine for which 30-second OEI and 2-minute OEI ratings 
are desired, that incorporates a means for automatic temperature control 
within its operating limitations in accordance with Sec. 33.28(k), must 
run for a period of 4 minutes at the maximum power-on rpm with the gas 
temperature at least 35 [deg]F (19 [deg]C) higher than the maximum 
operating limit at 30-second OEI rating. Following this run, the turbine 
assembly may exhibit distress beyond the limits for an overtemperature 
condition provided the engine is shown by analysis or test, as found 
necessary by the FAA, to maintain the integrity of the turbine assembly.
    (c) A separate test vehicle may be used for each test condition.

[Doc. No. 26019, 61 FR 31329, June 19, 1996, as amended by Amdt. 33-25, 
73 FR 48124, Aug. 18, 2008; Amdt. 33-26, 73 FR 48285, Aug. 19, 2008]



Sec. 33.89  Operation test.

    (a) The operation test must include testing found necessary by the 
Administrator to demonstrate--
    (1) Starting, idling, acceleration, overspeeding, ignition, 
functioning of the propeller (if the engine is designated to operate 
with a propeller);
    (2) Compliance with the engine response requirements of Sec. 33.73; 
and
    (3) The minimum power or thrust response time to 95 percent rated 
takeoff power or thrust, from power lever positions representative of 
minimum idle and of minimum flight idle, starting from stabilized idle 
operation, under the following engine load conditions:
    (i) No bleed air and power extraction for aircraft use.
    (ii) Maximum allowable bleed air and power extraction for aircraft 
use.
    (iii) An intermediate value for bleed air and power extraction 
representative of that which might be used as a maximum for aircraft 
during approach to a landing.
    (4) If testing facilities are not available, the determination of 
power extraction required in paragraph (a)(3)(ii) and (iii) of this 
section may be accomplished through appropriate analytical means.
    (b) The operation test must include all testing found necessary by 
the Administrator to demonstrate that the engine has safe operating 
characteristics throughout its specified operating envelope.

[Amdt. 33-4, 36 FR 5493, Mar. 24, 1971, as amended by Amdt. 33-6, 39 FR 
35469, Oct. 1, 1974; Amdt. 33-10, 49 FR 6853, Feb. 23, 1984]



Sec. 33.90  Initial maintenance inspection test.

    Each applicant, except an applicant for an engine being type 
certificated through amendment of an existing type certificate or 
through supplemental type certification procedures, must complete one of 
the following tests on an engine that substantially conforms to the type 
design to establish when the initial maintenance inspection is required:
    (a) An approved engine test that simulates the conditions in which 
the engine is expected to operate in service, including typical start-
stop cycles.
    (b) An approved engine test conducted in accordance with Sec. 
33.201 (c) through (f).

[Doc. No. FAA-2002-6717, 72 FR 1877, Jan. 16, 2007]



Sec. 33.91  Engine system and component tests.

    (a) For those systems or components that cannot be adequately 
substantiated in accordance with endurance testing of Sec. 33.87, the 
applicant must conduct additional tests to demonstrate that the systems 
or components are able to perform the intended functions in all declared 
environmental and operating conditions.
    (b) Temperature limits must be established for those components that 
require temperature controlling provisions in the aircraft installation 
to assure satisfactory functioning, reliability, and durability.
    (c) Each unpressurized hydraulic fluid tank may not fail or leak 
when subjected to a maximum operating temperature and an internal 
pressure of 5 p.s.i., and each pressurized hydraulic fluid tank must 
meet the requirements of Sec. 33.64.

[[Page 906]]

    (d) For an engine type certificated for use in supersonic aircraft, 
the systems, safety devices, and external components that may fail 
because of operation at maximum and minimum operating temperatures must 
be identified and tested at maximum and minimum operating temperatures 
and while temperature and other operating conditions are cycled between 
maximum and minimum operating values.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-6, 39 
FR 35469, Oct. 1, 1974; Amdt. 33-26, 73 FR 48285, Aug. 19, 2008; Amdt. 
33-27, 73 FR 55437, Sept. 25, 2008; Amdt. 33-27, 73 FR 57235, Oct. 2, 
2008]



Sec. 33.92  Rotor locking tests.

    If continued rotation is prevented by a means to lock the rotor(s), 
the engine must be subjected to a test that includes 25 operations of 
this means under the following conditions:
    (a) The engine must be shut down from rated maximum continuous 
thrust or power; and
    (b) The means for stopping and locking the rotor(s) must be operated 
as specified in the engine operating instructions while being subjected 
to the maximum torque that could result from continued flight in this 
condition; and
    (c) Following rotor locking, the rotor(s) must be held stationary 
under these conditions for five minutes for each of the 25 operations.

[Doc. No. 28107, 61 FR 28433, June 4, 1996]



Sec. 33.93  Teardown inspection.

    (a) After completing the endurance testing of Sec. 33.87 (b), (c), 
(d), (e), or (g) of this part, each engine must be completely 
disassembled, and
    (1) Each component having an adjustment setting and a functioning 
characteristic that can be established independent of installation on 
the engine must retain each setting and functioning characteristic 
within the limits that were established and recorded at the beginning of 
the test; and
    (2) Each engine part must conform to the type design and be eligible 
for incorporation into an engine for continued operation, in accordance 
with information submitted in compliance with Sec. 33.4.
    (b) After completing the endurance testing of Sec. 33.87(f), each 
engine must be completely disassembled, and
    (1) Each component having an adjustment setting and a functioning 
characteristic that can be established independent of installation on 
the engine must retain each setting and functioning characteristic 
within the limits that were established and recorded at the beginning of 
the test; and
    (2) Each engine may exhibit deterioration in excess of that 
permitted in paragraph (a)(2) of this section, including some engine 
parts or components that may be unsuitable for further use. The 
applicant must show by inspection, analysis, test, or by any combination 
thereof as found necessary by the FAA, that structural integrity of the 
engine is maintained; or
    (c) In lieu of compliance with paragraph (b) of this section, each 
engine for which the 30-second OEI and 2-minute OEI ratings are desired, 
may be subjected to the endurance testing of Sec. Sec. 33.87 (b), (c), 
(d), or (e) of this part, and followed by the testing of Sec. 33.87(f) 
without intervening disassembly and inspection. However, the engine must 
comply with paragraph (a) of this section after completing the endurance 
testing of Sec. 33.87(f).

[Doc. No. 26019, 61 FR 31329, June 19, 1996, as amended by Amdt. 33-25, 
73 FR 48124, Aug. 18, 2008]



Sec. 33.94  Blade containment and rotor unbalance tests.

    (a) Except as provided in paragraph (b) of this section, it must be 
demonstrated by engine tests that the engine is capable of containing 
damage without catching fire and without failure of its mounting 
attachments when operated for at least 15 seconds, unless the resulting 
engine damage induces a self shutdown, after each of the following 
events:
    (1) Failure of the most critical compressor or fan blade while 
operating at maximum permissible r.p.m. The blade failure must occur at 
the outermost retention groove or, for integrally-bladed rotor discs, at 
least 80 percent of the blade must fail.

[[Page 907]]

    (2) Failure of the most critical turbine blade while operating at 
maximum permissible r.p.m. The blade failure must occur at the outermost 
retention groove or, for integrally-bladed rotor discs, at least 80 
percent of the blade must fail. The most critical turbine blade must be 
determined by considering turbine blade weight and the strength of the 
adjacent turbine case at case temperatures and pressures associated with 
operation at maximum permissible r.p.m.
    (b) Analysis based on rig testing, component testing, or service 
experience may be substitute for one of the engine tests prescribed in 
paragraphs (a)(1) and (a)(2) of this section if--
    (1) That test, of the two prescribed, produces the least rotor 
unbalance; and
    (2) The analysis is shown to be equivalent to the test.

(Secs. 313(a), 601, and 603, Federal Aviation Act of 1958 (49 U.S.C. 
1354(a), 1421, and 1423); and 49 U.S.C. 106(g) Revised, Pub. L. 97-449, 
Jan. 12, 1983)

[Amdt. 33-10, 49 FR 6854, Feb. 23, 1984]



Sec. 33.95  Engine-propeller systems tests.

    If the engine is designed to operate with a propeller, the following 
tests must be made with a representative propeller installed by either 
including the tests in the endurance run or otherwise performing them in 
a manner acceptable to the Administrator:
    (a) Feathering operation: 25 cycles.
    (b) Negative torque and thrust system operation: 25 cycles from 
rated maximum continuous power.
    (c) Automatic decoupler operation: 25 cycles from rated maximum 
continuous power (if repeated decoupling and recoupling in service is 
the intended function of the device).
    (d) Reverse thrust operation: 175 cycles from the flight-idle 
position to full reverse and 25 cycles at rated maximum continuous power 
from full forward to full reverse thrust. At the end of each cycle the 
propeller must be operated in reverse pitch for a period of 30 seconds 
at the maximum rotational speed and power specified by the applicant for 
reverse pitch operation.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-3, 32 
FR 3737, Mar. 4, 1967]



Sec. 33.96  Engine tests in auxiliary power unit (APU) mode.

    If the engine is designed with a propeller brake which will allow 
the propeller to be brought to a stop while the gas generator portion of 
the engine remains in operation, and remain stopped during operation of 
the engine as an auxiliary power unit (``APU mode''), in addition to the 
requirements of Sec. 33.87, the applicant must conduct the following 
tests:
    (a) Ground locking: A total of 45 hours with the propeller brake 
engaged in a manner which clearly demonstrates its ability to function 
without adverse effects on the complete engine while the engine is 
operating in the APU mode under the maximum conditions of engine speed, 
torque, temperature, air bleed, and power extraction as specified by the 
applicant.
    (b) Dynamic braking: A total of 400 application-release cycles of 
brake engagements must be made in a manner which clearly demonstrates 
its ability to function without adverse effects on the complete engine 
under the maximum conditions of engine acceleration/deceleration rate, 
speed, torque, and temperature as specified by the applicant. The 
propeller must be stopped prior to brake release.
    (c) One hundred engine starts and stops with the propeller brake 
engaged.
    (d) The tests required by paragraphs (a), (b), and (c) of this 
section must be performed on the same engine, but this engine need not 
be the same engine used for the tests required by Sec. 33.87.
    (e) The tests required by paragraphs (a), (b), and (c) of this 
section must be followed by engine disassembly to the extent necessary 
to show compliance with the requirements of Sec. 33.93(a) and Sec. 
33.93(b).

[Amdt. 33-11, 51 FR 10346, Mar. 25, 1986]



Sec. 33.97  Thrust reversers.

    (a) If the engine incorporates a reverser, the endurance 
calibration, operation, and vibration tests prescribed in this subpart 
must be run with the reverser installed. In complying with this section, 
the power control lever must be moved from one extreme position to the 
other in not more than one second except, if regimes of control 
operations

[[Page 908]]

are incorporated necessitating scheduling of the power-control lever 
motion in going from one extreme position to the other, a longer period 
of time is acceptable but not more than three seconds. In addition, the 
test prescribed in paragraph (b) of this section must be made. This test 
may be scheduled as part of the endurance run.
    (b) 175 reversals must be made from flight-idle forward thrust to 
maximum reverse thrust and 25 reversals must be made from rated takeoff 
thrust to maximum reverse thrust. After each reversal the reverser must 
be operated at full reverse thrust for a period of one minute, except 
that, in the case of a reverser intended for use only as a braking means 
on the ground, the reverser need only be operated at full reverse thrust 
for 30 seconds.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-3, 32 
FR 3737, Mar. 4, 1967]



Sec. 33.99  General conduct of block tests.

    (a) Each applicant may, in making a block test, use separate engines 
of identical design and construction in the vibration, calibration, 
endurance, and operation tests, except that, if a separate engine is 
used for the endurance test it must be subjected to a calibration check 
before starting the endurance test.
    (b) Each applicant may service and make minor repairs to the engine 
during the block tests in accordance with the service and maintenance 
instructions submitted in compliance with Sec. 33.4. If the frequency 
of the service is excessive, or the number of stops due to engine 
malfunction is excessive, or a major repair, or replacement of a part is 
found necessary during the block tests or as the result of findings from 
the teardown inspection, the engine or its parts must be subjected to 
any additional tests the Administrator finds necessary.
    (c) Each applicant must furnish all testing facilities, including 
equipment and competent personnel, to conduct the block tests.

[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33-6, 39 
FR 35470, Oct. 1, 1974; Amdt. 33-9, 45 FR 60181, Sept. 11, 1980]



        Subpart G_Special Requirements: Turbine Aircraft Engines

    Source: Docket No. FAA-2002-6717, 72 FR 1877, Jan. 16, 2007, unless 
otherwise noted.



Sec. 33.201  Design and test requirements for Early ETOPS eligibility.

    An applicant seeking type design approval for an engine to be 
installed on a two-engine airplane approved for ETOPS without the 
service experience specified in part 25, appendix K, K25.2.1 of this 
chapter, must comply with the following:
    (a) The engine must be designed using a design quality process 
acceptable to the FAA, that ensures the design features of the engine 
minimize the occurrence of failures, malfunctions, defects, and 
maintenance errors that could result in an IFSD, loss of thrust control, 
or other power loss.
    (b) The design features of the engine must address problems shown to 
result in an IFSD, loss of thrust control, or other power loss in the 
applicant's other relevant type designs approved within the past 10 
years, to the extent that adequate service data is available within that 
10-year period. An applicant without adequate service data must show 
experience with and knowledge of problem mitigating design practices 
equivalent to that gained from actual service experience in a manner 
acceptable to the FAA.
    (c) Except as specified in paragraph (f) of this section, the 
applicant must conduct a simulated ETOPS mission cyclic endurance test 
in accordance with an approved test plan on an engine that substantially 
conforms to the type design. The test must:
    (1) Include a minimum of 3,000 representative service start-stop 
mission cycles and three simulated diversion cycles at maximum 
continuous thrust or power for the maximum diversion time for which 
ETOPS eligibility is sought. Each start-stop mission cycle must include 
the use of take-off, climb, cruise, descent, approach, and landing 
thrust or power and the use of thrust reverse (if applicable). The 
diversions must be evenly distributed over the duration of the test. The 
last diversion must be conducted within 100 cycles of the completion of 
the test.

[[Page 909]]

    (2) Be performed with the high speed and low speed main engine 
rotors independently unbalanced to obtain a minimum of 90 percent of the 
recommended field service maintenance vibration levels. For engines with 
three main engine rotors, the intermediate speed rotor must be 
independently unbalanced to obtain a minimum of 90 percent of the 
recommended production acceptance vibration level. The required peak 
vibration levels must be verified during a slow acceleration and 
deceleration run of the test engine covering the main engine rotor 
operating speed ranges.
    (3) Include a minimum of three million vibration cycles for each 60 
rpm incremental step of the typical high-speed rotor start-stop mission 
cycle. The test may be conducted using any rotor speed step increment 
from 60 to 200 rpm provided the test encompasses the typical service 
start-stop cycle speed range. For incremental steps greater than 60 rpm, 
the minimum number of vibration cycles must be linearly increased up to 
ten million cycles for a 200 rpm incremental step.
    (4) Include a minimum of 300,000 vibration cycles for each 60 rpm 
incremental step of the high-speed rotor approved operational speed 
range between minimum flight idle and cruise power not covered by 
paragraph (c)(3) of this section. The test may be conducted using any 
rotor speed step increment from 60 to 200 rpm provided the test 
encompasses the applicable speed range. For incremental steps greater 
than 60 rpm the minimum number of vibration cycles must be linearly 
increased up to 1 million for a 200 rpm incremental step.
    (5) Include vibration surveys at periodic intervals throughout the 
test. The equivalent value of the peak vibration level observed during 
the surveys must meet the minimum vibration requirement of Sec. 
33.201(c)(2).
    (d) Prior to the test required by paragraph (c) of this section, the 
engine must be subjected to a calibration test to document power and 
thrust characteristics.
    (e) At the conclusion of the testing required by paragraph (c) of 
this section, the engine must:
    (1) Be subjected to a calibration test at sea-level conditions. Any 
change in power or thrust characteristics must be within approved 
limits.
    (2) Be visually inspected in accordance with the on-wing inspection 
recommendations and limits contained in the Instructions for Continued 
Airworthiness submitted in compliance with Sec. 33.4.
    (3) Be completely disassembled and inspected--
    (i) In accordance with the applicable inspection recommendations and 
limits contained in the Instructions for Continued Airworthiness 
submitted in compliance with Sec. 33.4;
    (ii) With consideration of the causes of IFSD, loss of thrust 
control, or other power loss identified by paragraph (b) of this 
section; and
    (iii) In a manner to identify wear or distress conditions that could 
result in an IFSD, loss of thrust control, or other power loss not 
specifically identified by paragraph (b) of this section or addressed 
within the Instructions for Continued Airworthiness.
    (4) Not show wear or distress to the extent that could result in an 
IFSD, loss of thrust control, or other power loss within a period of 
operation before the component, assembly, or system would likely have 
been inspected or functionally tested for integrity while in service. 
Such wear or distress must have corrective action implemented through a 
design change, a change to maintenance instructions, or operational 
procedures before ETOPS eligibility is granted. The type and frequency 
of wear and distress that occurs during the engine test must be 
consistent with the type and frequency of wear and distress that would 
be expected to occur on ETOPS eligible engines.
    (f) An alternative mission cycle endurance test that provides an 
equivalent demonstration of the unbalance and vibration specified in 
paragraph (c) of this section may be used when approved by the FAA.
    (g) For an applicant using the simulated ETOPS mission cyclic 
endurance test to comply with Sec. 33.90, the test may be interrupted 
so that the engine may be inspected by an on-wing or other method, using 
criteria acceptable

[[Page 910]]

to the FAA, after completion of the test cycles required to comply with 
Sec. 33.90(a). Following the inspection, the ETOPS test must be resumed 
to complete the requirements of this section.



  Sec. Appendix A to Part 33--Instructions for Continued Airworthiness

                              a33.1 general

    (a) This appendix specifies requirements for the preparation of 
Instructions for Continued Airworthiness as required by Sec. 33.4.
    (b) The Instructions for Continued Airworthiness for each engine 
must include the Instructions for Continued Airworthiness for all engine 
parts. If Instructions for Continued Airworthiness are not supplied by 
the engine part manufacturer for an engine part, the Instructions for 
Continued Airworthiness for the engine must include the information 
essential to the continued airworthiness of the engine.
    (c) The applicant must submit to the FAA a program to show how 
changes to the Instructions for Continued Airworthiness made by the 
applicant or by the manufacturers of engine parts will be distributed.

                              a33.2 format

    (a) The Instructions for Continued Airworthiness must be in the form 
of a manual or manuals as appropriate for the quantity of data to be 
provided.
    (b) The format of the manual or manuals must provide for a practical 
arrangement.

                              a33.3 content

    The contents of the manual or manuals must be prepared in the 
English language. The Instructions for Continued Airworthiness must 
contain the following manuals or sections, as appropriate, and 
information:
    (a) Engine Maintenance Manual or Section. (1) Introduction 
information that includes an explanation of the engine's features and 
data to the extent necessary for maintenance or preventive maintenance.
    (2) A detailed description of the engine and its components, 
systems, and installations.
    (3) Installation instructions, including proper procedures for 
uncrating, deinhibiting, acceptance checking, lifting, and attaching 
accessories, with any necessary checks.
    (4) Basic control and operating information describing how the 
engine components, systems, and installations operate, and information 
describing the methods of starting, running, testing, and stopping the 
engine and its parts including any special procedures and limitations 
that apply.
    (5) Servicing information that covers details regarding servicing 
points, capacities of tanks, reservoirs, types of fluids to be used, 
pressures applicable to the various systems, locations of lubrication 
points, lubricants to be used, and equipment required for servicing.
    (6) Scheduling information for each part of the engine that provides 
the recommended periods at which it should be cleaned, inspected, 
adjusted, tested, and lubricated, and the degree of inspection the 
applicable wear tolerances, and work recommended at these periods. 
However, the applicant may refer to an accessory, instrument, or 
equipment manufacturer as the source of this information if the 
applicant shows that the item has an exceptionally high degree of 
complexity requiring specialized maintenance techniques, test equipment, 
or expertise. The recommended overhaul periods and necessary cross 
references to the Airworthiness Limitations section of the manual must 
also be included. In addition, the applicant must include an inspection 
program that includes the frequency and extent of the inspections 
necessary to provide for the continued airworthiness of the engine.
    (7) Troubleshooting information describing probable malfunctions, 
how to recognize those malfunctions, and the remedial action for those 
malfunctions.
    (8) Information describing the order and method of removing the 
engine and its parts and replacing parts, with any necessary precautions 
to be taken. Instructions for proper ground handling, crating, and 
shipping must also be included.
    (9) A list of the tools and equipment necessary for maintenance and 
directions as to their method of use.
    (b) Engine Overhaul Manual or Section. (1) Disassembly information 
including the order and method of disassembly for overhaul.
    (2) Cleaning and inspection instructions that cover the materials 
and apparatus to be used and methods and precautions to be taken during 
overhaul. Methods of overhaul inspection must also be included.
    (3) Details of all fits and clearances relevant to overhaul.
    (4) Details of repair methods for worn or otherwise substandard 
parts and components along with the information necessary to determine 
when replacement is necessary.
    (5) The order and method of assembly at overhaul.
    (6) Instructions for testing after overhaul.
    (7) Instructions for storage preparation, including any storage 
limits.
    (8) A list of tools needed for overhaul.
    (c) ETOPS Requirements. For an applicant seeking eligibility for an 
engine to be installed on an airplane approved for ETOPS, the 
Instructions for Continued Airworthiness must include procedures for 
engine condition monitoring. The engine condition monitoring procedures 
must be able to determine prior to flight, whether an engine is capable

[[Page 911]]

of providing, within approved engine operating limits, maximum 
continuous power or thrust, bleed air, and power extraction required for 
a relevant engine inoperative diversion. For an engine to be installed 
on a two-engine airplane approved for ETOPS, the engine condition 
monitoring procedures must be validated before ETOPS eligibility is 
granted.

                 A33.4 airworthiness limitations section

    The Instructions for Continued Airworthiness must contain a section 
titled Airworthiness Limitations that is segregated and clearly 
distinguishable from the rest of the manual.
    (a) For all engines:
    (1) The Airworthiness Limitations section must set forth each 
mandatory replacement time, inspection interval, and related procedure 
required for type certification. If the Instructions for Continued 
Airworthiness consist of multiple documents, the section required under 
this paragraph must be included in the principal manual.
    (2) This section must contain a legible statement in a prominent 
location that reads: ``The Airworthiness Limitations section is FAA 
approved and specifies maintenance required under Sec. Sec. 43.16 and 
91.403 of Title 14 of the Code of Federal Regulations unless an 
alternative program has been FAA approved.''
    (b) For rotorcraft engines having 30-second OEI and 2-minute OEI 
ratings:
    (1) The Airworthiness Limitations section must also prescribe the 
mandatory post-flight inspections and maintenance actions associated 
with any use of either 30-second OEI or 2-minute OEI ratings.
    (2) The applicant must validate the adequacy of the inspections and 
maintenance actions required under paragraph (b)(1) of this section 
A33.4.
    (3) The applicant must establish an in-service engine evaluation 
program to ensure the continued adequacy of the instructions for 
mandatory post-flight inspections and maintenance actions prescribed 
under paragraph (b)(1) of this section A33.4 and of the data for Sec. 
33.5(b)(4) pertaining to power availability. The program must include 
service engine tests or equivalent service engine test experience on 
engines of similar design and evaluations of service usage of the 30-
second OEI or 2-minute OEI ratings.

[Amdt. 33-9, 45 FR 60181, Sept. 11, 1980, as amended by Amdt. 33-13, 54 
FR 34330, Aug. 18, 1989; Amdt. 33-21, 72 FR 1878, Jan. 16, 2007; Amdt. 
33-25, 73 FR 48124, Aug. 18, 2008]



     Sec. Appendix B to Part 33--Certification Standard Atmospheric 
                     Concentrations of Rain and Hail

    Figure B1, Table B1, Table B2, Table B3, and Table B4 specify the 
atmospheric concentrations and size distributions of rain and hail for 
establishing certification, in accordance with the requirements of Sec. 
33.78(a)(2). In conducting tests, normally by spraying liquid water to 
simulate rain conditions and by delivering hail fabricated from ice to 
simulate hail conditions, the use of water droplets and hail having 
shapes, sizes and distributions of sizes other than those defined in 
this appendix B, or the use of a single size or shape for each water 
droplet or hail, can be accepted, provided that applicant shows that the 
substitution does not reduce the severity of the test.

[[Page 912]]

[GRAPHIC] [TIFF OMITTED] TR26MR98.000


    Table B1--Certification Standard Atmospheric Rain Concentrations
------------------------------------------------------------------------
                                                Rain water content (RWC)
               Altitude (feet)                   (grams water/meter \3\
                                                          air)
------------------------------------------------------------------------
0............................................                       20.0
20,000.......................................                       20.0
26,300.......................................                       15.2
32,700.......................................                       10.8
39,300.......................................                        7.7
46,000.......................................                        5.2
------------------------------------------------------------------------
RWC values at other altitudes may be determined by linear interpolation.
Note: Source of data--Results of the Aerospace Industries Association
  (AIA) Propulsion Committee Study, Project PC 338-1, June 1990.


    Table B2--Certification Standard Atmospheric Hail Concentrations
------------------------------------------------------------------------
                                                Hail water content (HWC)
               Altitude (feet)                   (grams water/meter \3\
                                                          air)
------------------------------------------------------------------------
0............................................                        6.0
7,300........................................                        8.9
8,500........................................                        9.4
10,000.......................................                        9.9
12,000.......................................                       10.0
15,000.......................................                       10.0
16,000.......................................                        8.9
17,700.......................................                        7.8
19,300.......................................                        6.6
21,500.......................................                        5.6
24,300.......................................                        4.4
29,000.......................................                        3.3
46,000.......................................                        0.2
------------------------------------------------------------------------
HWC values at other altitudes may be determined by linear interpolation.
  The hail threat below 7,300 feet and above 29,000 feet is based on
  linearly extrapolated data.
Note: Source of data--Results of the Aerospace Industries Association
  (AIA Propulsion Committee (PC) Study, Project PC 338-1, June 1990.


     Table B3--Certification Standard Atmospheric Rain Droplet Size
                              Distribution
------------------------------------------------------------------------
                                                 Contribution total RWC
          Rain droplet diameter (mm)                      (%)
------------------------------------------------------------------------
0-0.49.......................................                          0
0.50-0.99....................................                       2.25
1.00-1.49....................................                       8.75
1.50-1.99....................................                      16.25
2.00-2.49....................................                      19.00
2.50-2.99....................................                      17.75
3.00-3.49....................................                      13.50
3.50-3.99....................................                       9.50
4.00-4.49....................................                       6.00
4.50-4.99....................................                       3.00
5.00-5.49....................................                       2.00
5.50-5.99....................................                       1.25
6.00-6.49....................................                       0.50
6.50-7.00....................................                       0.25
                                              --------------------------
    Total....................................                     100.00
------------------------------------------------------------------------
Median diameter of rain droplets in 2.66 mm
Note: Source of data--Results of the Aerospace Industries Association
  (AIA Propulsion Committee (PC) Study, Project PC 338-1, June 1990.


   Table B4--Certification Standard Atmospheric Hail Size Distribution
------------------------------------------------------------------------
                                                 Contribution total HWC
              Hail diameter (mm)                          (%)
------------------------------------------------------------------------
0-4.9........................................                          0
5.0-9.9......................................                      17.00
10.0-14.9....................................                      25.00
15.0-19.9....................................                      22.50
20.0-24.9....................................                      16.00
25.0-29.9....................................                       9.75
30.0-34.9....................................                       4.75
35.0-39.9....................................                       2.50
40.0-44.9....................................                       1.50
45.0-49.9....................................                       0.75
50.0-55.0....................................                       0.25
                                              --------------------------

[[Page 913]]

 
    Total....................................                     100.00
------------------------------------------------------------------------
Median diameter of hail is 16 mm
Note: Source of data--Results of the Aerospace Industries Association
  (AIA Propulsion Committee (PC) Study, Project PC 338-1, June 1990.


[Doc. No. 28652, 63 FR 14799, Mar. 26, 1998]



                      Sec. Appendix C to Part 33--

                               [Reserved]

    Effective Date Note: At 79 FR 65538, Nov. 4, 2014, appendix C to 
part 33 was added and reserved, effective Jan. 5, 2015.



 Sec. Appendix D to Part 33--Mixed Phase and Ice Crystal Icing Envelope 
                        (Deep Convective Clouds)

    The ice crystal icing envelope is depicted in Figure D1 of this 
Appendix.
[GRAPHIC] [TIFF OMITTED] TR04NO14.008

    Within the envelope, total water content (TWC) in g/m\3\ has been 
determined based upon the adiabatic lapse defined by the convective rise 
of 90% relative humidity air from sea level to higher altitudes and 
scaled by a factor of 0.65 to a standard cloud length of 17.4 nautical 
miles. Figure D2 of this Appendix displays TWC for this distance over a 
range of ambient temperature within the boundaries of the ice crystal 
envelope specified in Figure D1 of this Appendix.

[[Page 914]]

[GRAPHIC] [TIFF OMITTED] TR04NO14.009

    Ice crystal size median mass dimension (MMD) range is 50-200 microns 
(equivalent spherical size) based upon measurements near convective 
storm cores.
    The TWC can be treated as completely glaciated (ice crystal) except 
as noted in the Table 1 of this Appendix.

               Table 1--Supercooled Liquid Portion of TWC
------------------------------------------------------------------------
                                          Horizontal cloud     LWC--  g/
      Temperature range--deg C         length--nautical miles     m\3\
------------------------------------------------------------------------
0 to -20............................  <=50...................      <=1.0
0 to -20............................  Indefinite.............      <=0.5
< -20...............................  .......................          0
------------------------------------------------------------------------

    The TWC levels displayed in Figure D2 of this Appendix represent TWC 
values for a standard exposure distance (horizontal cloud length) of 
17.4 nautical miles that must be adjusted with length of icing exposure.

[[Page 915]]

[GRAPHIC] [TIFF OMITTED] TR04NO14.010


[79 FR 65538, Nov. 4, 2014]

    Effective Date Note: At 79 FR 65538, Nov. 4, 2014, Part 33, Appendix 
D was added, effective Jan. 5, 2015.



PART 34_FUEL VENTING AND EXHAUST EMISSION REQUIREMENTS FOR TURBINE
ENGINE POWERED AIRPLANES--Table of Contents



                      Subpart A_General Provisions

Sec.
34.1 Definitions.
34.2 Abbreviations.
34.3 General requirements.
34.4 [Reserved]
34.5 Special test procedures.
34.6 Aircraft safety.
34.7 Exemptions.
34.9 Exceptions.

  Subpart B_Engine Fuel Venting Emissions (New and In-Use Aircraft Gas 
                            Turbine Engines)

34.10 Applicability.
34.11 Standard for fuel venting emissions.

     Subpart C_Exhaust Emissions (New Aircraft Gas Turbine Engines)

34.20 Applicability.
34.21 Standards for exhaust emissions.
34.23 Exhaust Emission Standards for Engines Manufactured on and after 
          July 18, 2012.

    Subpart D_Exhaust Emissions (In-Use Aircraft Gas Turbine Engines)

34.30 Applicability.
34.31 Standards for exhaust emissions.

                   Subpart E_Certification Provisions

34.48 Derivative engines for emissions certification purposes.

Subpart F [Reserved]

Subpart G_Test Procedures for Engine Exhaust Gaseous Emissions (Aircraft 
                    and Aircraft Gas Turbine Engines)

34.60 Introduction.
34.61-34.71 [Reserved]

    Authority: 42 U.S.C. 4321 et seq., 7572; 49 U.S.C. 106(g), 40113, 
44701-44702, 44704, 44714.

[[Page 916]]


    Source: Docket No. 25613, 55 FR 32861, Aug. 10, 1990, unless 
otherwise noted.



                      Subpart A_General Provisions



Sec. 34.1  Definitions.

    As used in this part, all terms not defined herein shall have the 
meaning given them in the Clean Air Act, as amended (42 U.S.C. 7401 et. 
seq.):
    Act means the Clean Air Act, as amended (42 U.S.C. 7401 et. seq.).
    Administrator means the Administrator of the Federal Aviation 
Administration or any person to whom he has delegated his authority in 
the matter concerned.
    Administrator of the EPA means the Administrator of the 
Environmental Protection Agency and any other officer or employee of the 
Environmental Protection Agency to whom the authority involved may be 
delegated.
    Aircraft as used in this part means any airplane as defined in 14 
CFR part 1 for which a U.S. standard airworthiness certificate or 
equivalent foreign airworthiness certificate is issued.
    Aircraft engine means a propulsion engine which is installed in, or 
which is manufactured for installation in, an aircraft.
    Aircraft gas turbine engine means a turboprop, turbofan, or turbojet 
aircraft engine.
    Characteristic level has the meaning given in Appendix 6 of ICAO 
Annex 16 as of July 2008. The characteristic level is a calculated 
emission level for each pollutant based on a statistical assessment of 
measured emissions from multiple tests.\1\
---------------------------------------------------------------------------

    \1\ This incorporation by reference was approved by the Director of 
the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 
51. This document can be obtained from the ICAO, Document Sales Unit, 
999 University Street, Montreal, Quebec H3C 5H7, Canada, phone +1 514-
954-8022, or www.icao.int or sales14icao.int. Copies can be reviewed at 
the FAA New England Regional Office, 12 New England Executive Park, 
Burlington, Massachusetts, 781-238-7101, or at the National Archives and 
Records Administration (NARA). For information on the availability of 
this material at NARA, call 202-741-6030, or go to: http://
www.archives.gov/federal--register/code--of--federal--regulations/ibr--
locations.html.
---------------------------------------------------------------------------

    Class TP means all aircraft turboprop engines.
    Class TF means all turbofan or turbojet aircraft engines or aircraft 
engines designed for applications that otherwise would have been 
fulfilled by turbojet and turbofan engines except engines of class T3, 
T8, and TSS.
    Class T3 means all aircraft gas turbine engines of the JT3D model 
family.
    Class T8 means all aircraft gas turbine engines of the JT8D model 
family.
    Class TSS means all aircraft gas turbine engines employed for 
propulsion of aircraft designed to operate at supersonic flight speeds.
    Commercial aircraft engine means any aircraft engine used or 
intended for use by an ``air carrier'' (including those engaged in 
``intrastate air transportation'') or a ``commercial operator'' 
(including those engaged in ``intrastate air transportation'') as these 
terms are defined in Title 49 of the United States Code and Title 14 of 
the Code of Federal Regulations.
    Commercial aircraft gas turbine engine means a turboprop, turbofan, 
or turbojet commercial aircraft engine.
    Date of manufacture of an engine is the date the inspection 
acceptance records reflect that the engine is complete and meets the FAA 
approved type design.
    Derivative engine for emissions certification purposes means an 
engine that has the same or similar emissions characteristics as an 
engine covered by a U.S. type certificate issued under 14 CFR part 33. 
These characteristics are specified in Sec. 34.48.
    Emission measurement system means all of the equipment necessary to 
transport the emission sample and measure the level of emissions. This 
includes the sample system and the instrumentation system.
    Engine model means all commercial aircraft turbine engines which are 
of the same general series, displacement, and design characteristics and 
are approved under the same type certificate.
    Excepted, as used in Sec. 34.9, means an engine that may be 
produced and sold that does not meet otherwise applicable standards. 
Excepted engines must conform to regulatory conditions specified for an 
exception in Sec. 34.9. Excepted engines are subject to the standards 
of

[[Page 917]]

this part even though they are not required to comply with the otherwise 
applicable requirements. Engines excepted with respect to certain 
standards must comply with other standards from which they are not 
specifically excepted.
    Exempt means an engine that does not meet certain applicable 
standards but may be produced and sold under the terms allowed by a 
grant of exemption issued pursuant to Sec. 34.7 of this part and part 
11 of this chapter. Exempted engines must conform to regulatory 
conditions specified in the exemption as well as other applicable 
regulations. Exempted engines are subject to the standards of this part 
even though they are not required to comply with the otherwise 
applicable requirements. Engines exempted with respect to certain 
standards must comply with other standards as a condition of the 
exemption.
    Exhaust emissions means substances emitted into the atmosphere from 
the exhaust discharge nozzle of an aircraft or aircraft engine.
    Fuel venting emissions means raw fuel, exclusive of hydrocarbons in 
the exhaust emissions, discharged from aircraft gas turbine engines 
during all normal ground and flight operations.
    In-use aircraft gas turbine engine means an aircraft gas turbine 
engine which is in service.
    Introduction date means the date of manufacture of the first 
individual production engine of a given engine model or engine type 
certificate family to be certificated. Neither test engines nor engines 
not placed into service affect this date.
    New aircraft turbine engine means an aircraft gas turbine engine 
which has never been in service.
    Power setting means the power or thrust output of an engine in terms 
of kilonewtons thrust for turbojet and turbofan engines or shaft power 
in terms of kilowatts for turboprop engines.
    Rated output (rO) means the maximum power/thrust available for 
takeoff at standard day conditions as approved for the engine by the 
Federal Aviation Administration, including reheat contribution where 
applicable, but excluding any contribution due to water injection, 
expressed in kilowatts or kilonewtons (as applicable), rounded to at 
least three significant figures.
    Rated pressure ratio (rPR) means the ratio between the combustor 
inlet pressure and the engine inlet pressure achieved by an engine 
operation at rated output, rounded to at least three significant 
figures.
    Reference day conditions means the reference ambient conditions to 
which the gaseous emissions (HC and smoke) are to be corrected. The 
reference day conditions are as follows: Temperature=15 [deg]C, specific 
humidity=0.00629 kg H2 O/kg of dry air, and pressure=101325 
Pa.
    Sample system means the system which provides for the transportation 
of the gaseous emission sample from the sample probe to the inlet of the 
instrumentation system.
    Shaft power means only the measured shaft power output of a 
turboprop engine.
    Smoke means the matter in exhaust emissions which obscures the 
transmission of light.
    Smoke number (SN) means the dimensionless term quantifying smoke 
emissions.
    Standard day conditions means the following ambient conditions: 
temperature = 15 [deg]C, specific humidity = 0.00634 kg H2O/
kg dry air, and pressure = 101.325 kPa.
    Taxi/idle (in) means those aircraft operations involving taxi and 
idle between the time of landing roll-out and final shutdown of all 
propulsion engines.
    Taxi/idle (out) means those aircraft operations involving taxi and 
idle between the time of initial starting of the propulsion engine(s) 
used for the taxi and the turn onto the duty runway.
    Tier, as used in this part, is a designation related to the 
NOX emission standard for the engine as specified in Sec. 
34.21 or Sec. 34.23 of this part (e.g., Tier 0).

[Doc. No. 25613, 55 FR 32861, Aug. 10, 1990; 55 FR 37287, Sept. 10, 
1990, as amended by Amdt. 34-3, 64 FR 5558, Feb. 3, 1999; Amdt. 34-5, 77 
FR 76849, Dec. 31, 2012; Amdt. 34-5A, 78 FR 63016, Oct. 23, 2013]

[[Page 918]]



Sec. 34.2  Abbreviations.

    The abbreviations used in this part have the following meanings in 
both upper and lower case:
CO2 Carbon dioxide
CO Carbon monoxide
EPA United States Environmental Protection Agency
FAA Federal Aviation Administration, United States Department of 
    Transportation
g Gram(s)
HC Hydrocarbon(s)
HP Horsepower
hr Hour(s)
H20 water
kg Kilogram(s)
kJ Kilojoule(s)
kN Kilonewton(s)
kW Kilowatt(s)
lb Pound(s)
LTO Landing and takeoff
min Minute(s)
NOX Oxides of nitrogen
Pa Pascal(s)
rO Rated output
rPR Rated pressure ratio
sec Second(s)
SP Shaft power
SN Smoke number
T Temperature, degrees Kelvin
TIM Time in mode
[deg]C Degrees Celsius
% Percent

[Doc. No. 25613, 55 FR 32861, Aug. 10, 1990, as amended by Amdt. 34-3, 
64 FR 5559, Feb. 3, 1999; Amdt. 34-5, 77 FR 76850, Dec. 31, 2012]



Sec. 34.3  General requirements.

    (a) This part provides for the approval or acceptance by the 
Administrator or the Administrator of the EPA of testing and sampling 
methods, analytical techniques, and related equipment not identical to 
those specified in this part. Before either approves or accepts any such 
alternate, equivalent, or otherwise nonidentical procedures or 
equipment, the Administrator or the Administrator of the EPA shall 
consult with the other in determining whether or not the action requires 
rulemaking under sections 231 and 232 of the Clean Air Act, as amended, 
consistent with the responsibilities of the Administrator of the EPA and 
the Secretary of Transportation under sections 231 and 232 of the Clean 
Air Act.
    (b) Under section 232 of the Act, the Secretary of Transportation 
issues regulations to ensure compliance with 40 CFR part 87. This 
authority has been delegated to the Administrator of the FAA (49 CFR 
1.47).
    (c) U.S. airplanes. This part applies to civil airplanes that are 
powered by aircraft gas turbine engines of the classes specified herein 
and that have U.S. standard airworthiness certificates.
    (d) Foreign airplanes. Pursuant to the definition of ``aircraft'' in 
40 CFR 87.1, this regulation applies to civil airplanes that are powered 
by aircraft gas turbine engines of the classes specified herein and that 
have foreign airworthiness certificates that are equivalent to U.S. 
standard airworthiness certificates. This regulation applies only to 
those foreign civil airplanes that, if registered in the United States, 
would be required by applicable regulations to have a U.S. standard 
airworthiness certificate in order to conduct the operations intended 
for the airplane. Pursuant to 40 CFR 87.3(c), this regulation does not 
apply where it would be inconsistent with an obligation assumed by the 
United States to a foreign country in a treaty, convention, or 
agreement.
    (e) Reference in this regulation to 40 CFR part 87 refers to title 
40 of the Code of Federal Regulations, chapter I--Environmental 
Protection Agency, part 87, Control of Air Pollution from Aircraft and 
Aircraft Engines (40 CFR part 87).
    (f) This part contains regulations to ensure compliance with certain 
standards contained in 40 CFR part 87. If EPA takes any action, 
including the issuance of an exemption or issuance of a revised or 
alternate procedure, test method, or other regulation, the effect of 
which is to relax or delay the effective date of any provision of 40 CFR 
part 87 that is made applicable to an aircraft under this FAR, the 
Administrator of FAA will grant a general administrative waiver of its 
more stringent requirements until this FAR is amended to reflect the 
more relaxed requirements prescribed by EPA.
    (g) Unless otherwise stated, all terminology and abbreviations in 
this FAR that are defined in 40 CFR part 87 have the meaning specified 
in that part, and all terms in 40 CFR part 87

[[Page 919]]

that are not defined in that part but that are used in this FAR have the 
meaning given them in the Clean Air Act, as amended by Public Law 91-
604.
    (h) All interpretations of 40 CFR part 87 that are rendered by the 
EPA also apply to this FAR.
    (i) If the EPA, under 40 CFR 87.3(a), approves or accepts any 
testing and sampling procedures or methods, analytical techniques, or 
related equipment not identical to those specified in that part, this 
FAR requires an applicant to show that such alternate, equivalent, or 
otherwise nonidentical procedures have been complied with, and that such 
alternate equipment was used to show compliance, unless the applicant 
elects to comply with those procedures, methods, techniques, and 
equipment specified in 40 CFR part 87.
    (j) If the EPA, under 40 CFR 87.5, prescribes special test 
procedures for any aircraft or aircraft engine that is not susceptible 
to satisfactory testing by the procedures in 40 CFR part 87, the 
applicant must show the Administrator that those special test procedures 
have been complied with.
    (k) Wherever 40 CFR part 87 requires agreement, acceptance, or 
approval by the Administrator of the EPA, this FAR requires a showing 
that such agreement or approval has been obtained.
    (l) Pursuant to 42 U.S.C. 7573, no state or political subdivision 
thereof may adopt or attempt to enforce any standard respecting 
emissions of any air pollutant from any aircraft or engine thereof 
unless that standard is identical to a standard made applicable to the 
aircraft by the terms of this FAR.
    (m) If EPA, by regulation or exemption, relaxes a provision of 40 
CFR part 87 that is implemented in this FAR, no state or political 
subdivision thereof may adopt or attempt to enforce the terms of this 
FAR that are superseded by the relaxed requirement.
    (n) If any provision of this FAR is rendered inapplicable to a 
foreign aircraft as provided in 40 CFR 87.3(c) (international 
agreements), and Sec. 34.3(d) of this FAR, that provision may not be 
adopted or enforced against that foreign aircraft by a state or 
political subdivision thereof.
    (o) For exhaust emissions requirements of this FAR that apply 
beginning February 1, 1974, January 1, 1976, January 1, 1978, January 1, 
1984, and August 9, 1985, continued compliance with those requirements 
is shown for engines for which the type design has been shown to meet 
those requirements, if the engine is maintained in accordance with 
applicable maintenance requirements for 14 CFR chapter I. All methods of 
demonstrating compliance and all model designations previously found 
acceptable to the Administrator shall be deemed to continue to be an 
acceptable demonstration of compliance with the specific standards for 
which they were approved.
    (p) Each applicant must allow the Administrator to make, or witness, 
any test necessary to determine compliance with the applicable 
provisions of this FAR.

[Doc. No. 25613, 55 FR 32861, Aug. 10, 1990; 55 FR 37287, Sept. 10, 
1990; Amdt. 34-5, 77 FR 76850, Dec. 31, 2012]



Sec. 34.4  [Reserved]



Sec. 34.5  Special test procedures.

    The Administrator or the Administrator of the EPA may, upon written 
application by a manufacturer or operator of aircraft or aircraft 
engines, approve test procedures for any aircraft or aircraft engine 
that is not susceptible to satisfactory testing by the procedures set 
forth herein. Prior to taking action on any such application, the 
Administrator or the Administrator of the EPA shall consult with the 
other.



Sec. 34.6  Aircraft safety.

    (a) The provisions of this part will be revised if at any time the 
Administrator determines that an emission standard cannot be met within 
the specified time without creating a safety hazard.
    (b) Consistent with 40 CFR 87.6, if the FAA Administrator determines 
that any emission control regulation in this part cannot be safely 
applied to an aircraft, that provision may not be adopted or enforced 
against that aircraft by any state or political subdivision thereof.

[[Page 920]]



Sec. 34.7  Exemptions.

    Notwithstanding part 11 of the Federal Aviation Regulations (14 CFR 
part 11), all petitions for rulemaking involving either the substance of 
an emission standard or test procedure prescribed by the EPA that is 
incorporated in this FAR, or the compliance date for such standard or 
procedure, must be submitted to the EPA. Information copies of such 
petitions are invited by the FAA. Petitions for rulemaking or exemption 
involving provisions of this FAR that do not affect the substance or the 
compliance date of an emission standard or test procedure that is 
prescribed by the EPA, and petitions for exemptions under the provisions 
for which the EPA has specifically granted exemption authority to the 
Secretary of Transportation are subject to part 11 of the Federal 
Aviation Regulations (14 CFR part 11). Petitions for rulemaking or 
exemptions involving these FARs must be submitted to the FAA.
    (a) Exemptions based on flights for short durations at infrequent 
intervals. The emission standards of this part do not apply to engines 
which power aircraft operated in the United States for short durations 
at infrequent intervals. Such operations are limited to:
    (1) Flights of an aircraft for the purpose of export to a foreign 
country, including any flights essential to demonstrate the integrity of 
an aircraft prior to a flight to a point outside the United States.
    (2) Flights to a base where repairs, alterations or maintenance are 
to be performed, or to a point of storage, or for the purpose of 
returning an aircraft to service.
    (3) Official visits by representatives of foreign governments.
    (4) Other flights the Administrator determines, after consultation 
with the Administrator of the EPA, to be for short durations at 
infrequent intervals. A request for such a determination shall be made 
before the flight takes place.
    (b) Exemptions for very low production engine models. The emissions 
standards of this part do not apply to engines of very low production 
after the date of applicability. For the purpose of this part, ``very 
low production'' is limited to a maximum total production for United 
States civil aviation applications of no more than 200 units covered by 
the same type certificate after January 1, 1984. Engines manufactured 
under this provision must be reported to the FAA by serial number on or 
before the date of manufacture and exemptions granted under this 
provision are not transferable to any other engine. This exemption is 
limited to the requirements of Sec. 34.21 only.
    (c) Exemptions for new engines in other categories. The emissions 
standards of this part do not apply to engines for which the 
Administrator determines, with the concurrence of the Administrator of 
the EPA, that application of any standard under Sec. 34.21 is not 
justified, based upon consideration of--
    (1) Adverse economic impact on the manufacturer;
    (2) Adverse economic impact on the aircraft and airline industries 
at large;
    (3) Equity in administering the standards among all economically 
competing parties;
    (4) Public health and welfare effects; and
    (5) Other factors which the Administrator, after consultation with 
the Administrator of the EPA, may deem relevant to the case in question.
    (d) Applicants seeking exemption from other emissions standards of 
this part and 40 CFR part 87. Applicants must request exemption from 
both the FAA and the EPA, even where the underlying regulatory 
requirements are the same. The FAA and EPA will jointly consider such 
exemption requests, and will assure consistency in the respective agency 
determinations.
    (e) Applications for exemption from this part shall be submitted in 
duplicate to the Administrator in accordance with the procedures 
established by the Administrator in part 11.
    (f) The Administrator shall publish in the Federal Register the name 
of the organization to whom exemptions are granted and the period of 
such exemptions.
    (g) No state or political subdivision thereof may attempt to enforce 
a standard respecting emissions from an

[[Page 921]]

aircraft or engine if such aircraft or engine has been exempted from 
such standard under this part.

[Doc. No. 25613, 55 FR 32861, Aug. 10, 1990, as amended by Amdt. 34-5, 
77 FR 76850, Dec. 31, 2012]



Sec. 34.9  Exceptions.

    (a) Spare engines. Certain engines that meet the following 
description are excepted:
    (1) This exception allows production of an engine for installation 
on an in-service aircraft. A spare engine may not be installed on a new 
aircraft.
    (2) Each spare engine must be identical to a sub-model previously 
certificated to meet all applicable requirements.
    (3) A spare engine may be used only when the emissions of the spare 
do not exceed the certification requirements of the original engine, for 
all regulated pollutants.
    (4) No separate approval is required to produce spare engines.
    (5) The record for each engine excepted under this paragraph (c) 
must indicate that the engine was produced as an excepted spare engine.
    (6) Engines produced under this exception must be labeled ``EXCEPTED 
SPARE'' in accordance with Sec. 45.13 of this chapter.
    (b) On and after July 18, 2012, and before August 31, 2013, a 
manufacturer may produce up to six Tier 4 compliant engines that meet 
the NOX standards of paragraph (d)(1)(vi) of this section 
rather than Sec. 34.23(a)(2). No separate approval is required to 
produce these engines. Engines produced under this exception are to be 
labeled ``COMPLY'' in accordance with Sec. 45.13 of this chapter.

[Doc. No. FAA-2012-1333, 77 FR 76850, Dec. 31, 2012]



  Subpart B_Engine Fuel Venting Emissions (New and In-Use Aircraft Gas 
                            Turbine Engines)



Sec. 34.10  Applicability.

    (a) The provisions of this subpart are applicable to all new 
aircraft gas turbine engines of classes T3, T8, TSS, and TF equal to or 
greater than 36 kN (8,090 lb) rated output, manufactured on or after 
January 1, 1974, and to all in-use aircraft gas turbine engines of 
classes T3, T8, TSS, and TF equal to or greater than 36 kN (8,090 lb) 
rated output manufactured after February 1, 1974.
    (b) The provisions of this subpart are also applicable to all new 
aircraft gas turbine engines of class TF less than 36 kN (8,090 lb) 
rated output and class TP manufactured on or after January 1, 1975, and 
to all in-use aircraft gas turbine engines of class TF less than 36 kN 
(8,090 lb) rated output and class TP manufactured after January 1, 1975.

[Doc. No. FAA-2012-1333, 77 FR 76850, Dec. 31, 2012]



Sec. 34.11  Standard for fuel venting emissions.

    (a) No fuel venting emissions shall be discharged into the 
atmosphere from any new or in-use aircraft gas turbine engine subject to 
the subpart. This paragraph is directed at the elimination of 
intentional discharge to the atmosphere of fuel drained from fuel nozzle 
manifolds after engines are shut down and does not apply to normal fuel 
seepage from shaft seals, joints, and fittings.
    (b) Conformity with the standard set forth in paragraph (a) of this 
section shall be determined by inspection of the method designed to 
eliminate these emissions.
    (c) As applied to an airframe or an engine, any manufacturer or 
operator may show compliance with the fuel venting and emissions 
requirements of this section that were effective beginning February 1, 
1974 or January 1, 1975, by any means that prevents the intentional 
discharge of fuel from fuel nozzle manifolds after the engines are shut 
down. Acceptable means of compliance include one of the following:
    (1) Incorporation of an FAA-approved system that recirculates the 
fuel back into the fuel system.
    (2) Capping or securing the pressurization and drain valve.
    (3) Manually draining the fuel from a holding tank into a container.

[[Page 922]]



     Subpart C_Exhaust Emissions (New Aircraft Gas Turbine Engines)



Sec. 34.20  Applicability.

    The provisions of this subpart are applicable to all aircraft gas 
turbine engines of the classes specified beginning on the dates 
specified in Sec. 34.21.



Sec. 34.21  Standards for exhaust emissions.

    (a) Exhaust emissions of smoke from each new aircraft gas turbine 
engine of class T8 manufactured on or after February 1, 1974, shall not 
exceed a smoke number (SN) of 30.
    (b) Exhaust emissions of smoke from each new aircraft gas turbine 
engine of class TF and of rated output of 129 kN (29,000 lb) thrust or 
greater, manufactured on or after January 1, 1976, shall not exceed

SN = 83.6 (rO) -0.274 (rO is in kN).

    (c) Exhaust emission of smoke from each new aircraft gas turbine 
engine of class T3 manufactured on or after January 1, 1978, shall not 
exceed a smoke number (SN) of 25.
    (d) Gaseous exhaust emissions from each new aircraft gas turbine 
engine shall not exceed:
    (1) For Classes TF, T3, T8 engines greater than 26.7 kN (6,000 lb) 
rated output:
    (i) Engines manufactured on or after January 1, 1984:

Hydrocarbons: 19.6 g/kN rO.

    (ii) Engines manufactured on or after July 7, 1997:

Carbon Monoxide: 118 g/kN rO.

    (iii) Engines of a type or model of which the date of manufacture of 
the first individual production model was on or before December 31, 
1995, and for which the date of manufacture of the individual engine was 
on or before December 31, 1999 (Tier 2):

Oxides of Nitrogen: (40+2(rPR)) g/kN rO.

    (iv) Engines of a type or model of which the date of manufacture of 
the first individual production model was after December 31, 1995, or 
for which the date of manufacture of the individual engine was after 
December 31, 1999 (Tier 2):

Oxides of Nitrogen: (32+1.6(rPR)) g/kN rO.

    (v) The emission standards prescribed in paragraphs (d)(1)(iii) and 
(iv) of this section apply as prescribed beginning July 7, 1997.
    (vi) The emission standards of this paragraph apply as prescribed 
after December 18, 2005. For engines of a type or model of which the 
first individual production model was manufactured after December 31, 
2003 (Tier 4):
    (A) That have a rated pressure ratio of 30 or less and a maximum 
rated output greater than 89 kN:

Oxides of Nitrogen: (19 + 1.6(rPR)) g/kN rO.

    (B) That have a rated pressure ratio of 30 or less and a maximum 
rated output greater than 26.7 kN but not greater than 89 kN:

Oxides of Nitrogen: (37.572 + 1.6(rPR) - 0.2087(rO)) g/kN rO.

    (C) That have a rated pressure ratio greater than 30 but less than 
62.5, and a maximum rated output greater than 89 kN:

Oxides of Nitrogen: (7 + 2(rPR)) g/kN rO.

    (D) That have a rated pressure ratio greater than 30 but less than 
62.5, and a maximum rated output greater than 26.7 kN but not greater 
than 89 kN:

Oxides of Nitrogen: (42.71 + 1.4286(rPR) - 0.4013(rO) + 0.00642(rPR x 
rO)) g/kN rO.

    (E) That have a rated pressure ratio of 62.5 or more:

Oxides of Nitrogen: (32 + 1.6(rPR)) g/kN rO.

    (2) For Class TSS Engines manufactured on or after January 1, 1984:

Hydrocarbons: 140 (0.92)\rPR\ g/kN rO.

    (e) Smoke exhaust emissions from each gas turbine engine of the 
classes specified below shall not exceed:
    (1) For Class TF of rated output less than 26.7 kN (6,000 lb) 
manufactured on or after August 9, 1985:

SN = 83.6(rO) -0.274 (rO is in kN) not to exceed a maximum of 
SN = 50.


[[Page 923]]


    (2) For Classes T3, T8, TSS, and TF of rated output equal to or 
greater than 26.7 kN (6,000 lb) manufactured on or after January 1, 
1984:

SN = 83.6(rO) -0.274 (rO is in kN) not to exceed a maximum of 
SN = 50.

    (3) For Class TP of rated output equal to or greater than 1,000 kW 
manufactured on or after January 1, 1984:

SN = 187(rO) -0.168 (rO is in kW).

    (f) The standards set forth in paragraphs (a), (b), (c), (d), and 
(e) of this section refer to a composite gaseous emission sample 
representing the operation cycles and exhaust smoke emission emitted 
during operation of the engine as specified in the applicable sections 
of subpart G of this part, and measured and calculated in accordance 
with the procedures set forth in subpart G.
    (g) Where a gaseous emission standard is specified by a formula, 
calculate and round the standard to three significant figures or to the 
nearest 0.1 g/kN (for standards at or above 100 g/kN). Where a smoke 
standard is specified by a formula, calculate and round the standard to 
the nearest 0.1 SN. Engines comply with an applicable standard if the 
testing results show that the engine type certificate family's 
characteristic level does not exceed the numerical level of that 
standard, as described in Sec. 34.60.

[Doc. No. 25613, 55 FR 32861, Aug. 10, 1990; 55 FR 37287, Sept. 10, 
1990, as amended by Amdt. 34-3, 64 FR 5559, Feb. 3, 1999; Amdt. 34-4, 74 
FR 19127, Apr. 28, 2009; Amdt. 34-5, 77 FR 76851, Dec. 31, 2012]



Sec. 34.23  Exhaust Emission Standards for Engines Manufactured on
and after July 18, 2012.

    The standards of this section apply to aircraft engines manufactured 
on and after July 18, 2012, unless otherwise exempted or excepted. Where 
a gaseous emission standard is specified by a formula, calculate and 
round the standard to three significant figures or to the nearest 0.1 g/
kN (for standards at or above 100 g/kN). Where a smoke standard is 
specified by a formula, calculate and round the standard to the nearest 
0.1 SN. Engines comply with an applicable standard if the testing 
results show that the engine type certificate family's characteristic 
level does not exceed the numerical level of that standard, as described 
in Sec. 34.60.
    (a) Gaseous exhaust emissions from each new aircraft gas turbine 
engine shall not exceed:
    (1) For Classes TF, T3 and T8 of rated output less than 26.7 kN 
(6,000 lb) manufactured on and after July 18, 2012:

SN = 83.6(rO) -0.274 or 50.0, whichever is smaller

    (2) Except as provided in Sec. Sec. 34.9(b) and 34.21(c), for 
Classes TF, T3 and T8 engines manufactured on and after July 18, 2012, 
and for which the first individual production model was manufactured on 
or before December 31, 2013 (Tier 6):

                        Tier 6 Oxides of Nitrogen Emission Standards for Subsonic Engines
----------------------------------------------------------------------------------------------------------------
                                     Rated pressure ratio--
               Class                           rPR            Rated output rO (kN)            NOX (g/kN)
----------------------------------------------------------------------------------------------------------------
TF, T3, T8.........................  rPR <= 30.............  26.7 < rO <= 89.0.....  38.5486 + 1.6823 (rPR) -
                                                                                      0.2453 (rO) - (0.00308
                                                                                      (rPR) (rO)).
                                                            ----------------------------------------------------
                                                             rO  89.0...  16.72 + 1.4080 (rPR).
                                    ----------------------------------------------------------------------------
                                     30 < rPR < 82.6.......  26.7 < rO <= 89.0.....  46.1600 + 1.4286 (rPR) -
                                                                                      0.5303 (rO) + (0.00642
                                                                                      (rPR) (rO)).
                                    ----------------------------------------------------------------------------
                                                             rO  89.0...  -1.04 + 2.0 (rPR).
                                                            ----------------------------------------------------
                                     rPR = 82.6.  rO = 26.7..  32 + 1.6 (rPR).
----------------------------------------------------------------------------------------------------------------

    (3) Engines exempted from paragraph (a)(2) of this section produced 
on or before December 31, 2016 must be labeled ``EXEMPT NEW'' in 
accordance with

[[Page 924]]

Sec. 45.13 of this chapter. No exemptions to the requirements of 
paragraph (a)(2) of this section will be granted after December 31, 
2016.
    (4) For Class TSS Engines manufactured on and after July 18, 2012:

                                Gaseous Emission Standards for Supersonic Engines
----------------------------------------------------------------------------------------------------------------
                                   Rated output rO
             Class                    \1\ (kN)                 NOX (g/kN)                     CO (g/kN)
----------------------------------------------------------------------------------------------------------------
TSS............................  All...............  36 + 2.42 (rPR)..............  4,550 (rPR) -1.03
----------------------------------------------------------------------------------------------------------------
\1\ rO is the rated output with afterburning applied.

    (b) Gaseous exhaust emissions from each new aircraft gas turbine 
engine shall not exceed:
    (1) For Classes TF, T3 and T8 engines of a type or model of which 
the first individual production model was manufactured after December 
31, 2013 (Tier 8):

                        Tier 8 Oxides of Nitrogen Emission Standards for Subsonic Engines
----------------------------------------------------------------------------------------------------------------
                                     Rated pressure ratio--
               Class                           rPR            Rated output rO (kN)            NOX (g/kN)
----------------------------------------------------------------------------------------------------------------
TF, T3, T8.........................  rPR <= 30.............  26.7 < rO <= 89.0.....  40.052 + 1.5681 (rPR) -
                                                                                      0.3615 (rO) - (0.0018
                                                                                      (rPR) (rO)).
                                                            ----------------------------------------------------
                                                             rO  89.0...  7.88 + 1.4080 (rPR).
                                    ----------------------------------------------------------------------------
                                     30 < rPR < 104.7......  26.7 < rO <= 89.0.....  41.9435 + 1.505 (rPR) -
                                                                                      0.5823 (rO) + (0.005562
                                                                                      (rPR) (rO)).
                                                            ----------------------------------------------------
                                                             rO  89.0...  -9.88 + 2.0 (rPR).
                                    ----------------------------------------------------------------------------
                                     rPR = 104.7  rO = 26.7..  32 + 1.6 (rPR).
----------------------------------------------------------------------------------------------------------------

    (c) Engines (including engines that are determined to be derivative 
engines for the purposes of emission certification) type certificated 
with characteristic levels at or below the NOX standards of 
Sec. 34.21(d)(1)(vi) of this part (as applicable based on rated output 
and rated pressure ratio) and introduced before July 18, 2012, may be 
produced through December 31, 2012, without meeting the NOX 
standard of paragraph (a)(2) of this section.

[Doc. No. 34-5, 77 FR 76851, Dec. 31, 2012, as amended by Amdt. 34-5A, 
78 FR 63017, Oct. 23, 2013; 78 FR 65554, Nov. 1, 2013]



    Subpart D_Exhaust Emissions (In-use Aircraft Gas Turbine Engines)



Sec. 34.30  Applicability.

    The provisions of this subpart are applicable to all in-use aircraft 
gas turbine engines certificated for operation within the United States 
of the classes specified, beginning on the dates specified in Sec. 
34.31.



Sec. 34.31  Standards for exhaust emissions.

    (a) Exhaust emissions of smoke from each in-use aircraft gas turbine 
engine of Class T8, beginning February 1, 1974, shall not exceed a smoke 
number (SN) of 30.
    (b) Exhaust emissions of smoke from each in-use aircraft gas turbine 
engine of Class TF and of rated output of 129 kN (29,000 lb) thrust or 
greater, beginning January l, 1976, shall not exceed

SN=83.6(rO) -0.274 (rO is in kN).

    (c) The standards set forth in paragraphs (a) and (b) of this 
section refer to exhaust smoke emission emitted during operation of the 
engine as specified in the applicable sections of subpart G of this 
part, and measured and calculated in accordance with the procedures set 
forth in subpart G.

[Doc. No. FAA-2012-1333, 77 FR 76852, Dec. 31, 2012]

[[Page 925]]



                   Subpart E_Certification Provisions



Sec. 34.48  Derivative engines for emissions certification purposes.

    (a) General. A derivative engine for emissions certification 
purposes is an engine configuration that is determined to be similar in 
design to a previously certificated (original) engine for purposes of 
compliance with exhaust emissions standards (gaseous and smoke). A type 
certificate holder may request from the FAA a determination that an 
engine configuration is considered a derivative engine for emissions 
certification purposes. To be considered a derivative engine for 
emission purposes under this part, the configuration must have been 
derived from the original engine that was certificated to the 
requirements of part 33 of this chapter and one of the following:
    (1) The FAA has determined that a safety issue exists that requires 
an engine modification.
    (2) Emissions from the derivative engines are determined to be 
similar. In general, this means the emissions must meet the criteria 
specified in paragraph (b) of this section. The FAA may amend the 
criteria of paragraph (b) in unusual circumstances, for individual 
cases, consistent with good engineering judgment.
    (3) All of the regulated emissions from the derivative engine are 
lower than the original engine.
    (b) Emissions similarity. (1) The type certificate holder must 
demonstrate that the proposed derivative engine model's emissions meet 
the applicable standards and differ from the original model's emission 
rates only within the following ranges:
    (i) 3.0 g/kN for NOX.
    (ii) 1.0 g/kN for HC.
    (iii) 5.0 g/kN for CO.
    (iv) 2.0 SN for smoke.
    (2) If the characteristic level of the original certificated engine 
model (or any other sub-models within the emission type certificate 
family tested for certification) before modification is at or above 95% 
of the applicable standard for any pollutant, an applicant must measure 
the proposed derivative engine model's emissions for all pollutants to 
demonstrate that the derivative engine's resulting characteristic levels 
will not exceed the applicable emission standards. If the characteristic 
levels of the originally certificated engine model (and all other sub-
models within the emission type certificate family tested for 
certification) are below 95% of the applicable standard for each 
pollutant, the applicant may use engineering analysis consistent with 
good engineering judgment to demonstrate that the derivative engine will 
not exceed the applicable emission standards. The engineering analysis 
must address all modifications from the original engine, including those 
approved for previous derivative engines.
    (c) Continued production allowance. Derivative engines for emissions 
certification purposes may continue to be produced after the 
applicability date for new emissions standards when the engines conform 
to the specifications of this section.
    (d) Non-derivative engines. If the FAA determines that an engine 
model does not meet the requirements for a derivative engine for 
emissions certification purposes, the type certificate holder is 
required to demonstrate that the engine complies with the emissions 
standards applicable to a new engine type.

[Doc. No. 34-5, 77 FR 76852, Dec. 31, 2012]

Subpart F [Reserved]



Subpart G_Test Procedures for Engine Exhaust Gaseous Emissions (Aircraft 
                    and Aircraft Gas Turbine Engines)



Sec. 34.60  Introduction.

    (a) Use the equipment and procedures specified in Appendix 3, 
Appendix 5, and Appendix 6 of ICAO Annex 16, as applicable, to 
demonstrate whether engines meet the applicable gaseous emission 
standards specified in subpart C of this part. Measure the emissions of 
all regulated gaseous pollutants. Use the equipment and procedures 
specified in Appendix 2 and Appendix 6 of ICAO Annex 16 to determine 
whether engines meet the applicable smoke standard specified in subpart 
C of this part. The compliance demonstration consists of establishing a 
mean value from testing the specified number of engines, then

[[Page 926]]

calculating a ``characteristic level'' by applying a set of statistical 
factors that take into account the number of engines tested. Round each 
characteristic level to the same number of decimal places as the 
corresponding emission standard. For turboprop engines, use the 
procedures specified for turbofan engines, consistent with good 
engineering judgment.
    (b) Use a test fuel that meets the specifications described in 
Appendix 4 of ICAO Annex 16. The test fuel must not have additives whose 
purpose is to suppress smoke, such as organometallic compounds.
    (c) Prepare test engines by including accessories that are available 
with production engines if they can reasonably be expected to influence 
emissions. The test engine may not extract shaft power or bleed service 
air to provide power to auxiliary gearbox-mounted components required to 
drive aircraft systems.
    (d) Test engines must reach a steady operating temperature before 
the start of emission measurements.
    (e) In consultation with the EPA, the FAA may approve alternative 
procedures for measuring emissions, including testing and sampling 
methods, analytical techniques, and equipment specifications that differ 
from those specified in this part. Manufacturers and operators may 
request approval of alternative procedures by written request with 
supporting justification to the FAA Aircraft Certification Office and to 
the Designated EPA Program Officer. To be approved, one of the following 
conditions must be met:
    (1) The engine cannot be tested using the specified procedures; or
    (2) The alternative procedure is shown to be equivalent to, or more 
accurate or precise than, the specified procedure.
    (f) The following landing and takeoff (LTO) cycles apply for 
emissions testing and for calculating weighted LTO values:

                                        LTO Test Cycles and Time in Mode
----------------------------------------------------------------------------------------------------------------
                                                                        Class
                                   -----------------------------------------------------------------------------
               Mode                            TP                    TF, T3, T8                    TSS
                                   -----------------------------------------------------------------------------
                                     TIM (min)     % of rO     TIM (min)     % of rO     TIM (min)     % of rO
----------------------------------------------------------------------------------------------------------------
Taxi/idle.........................         26.0            7         26.0            7         26.0          5.8
Takeoff...........................          0.5          100          0.7          100          1.2          100
Climbout..........................          2.5           90          2.2           85          2.0           65
Descent...........................           NA           NA           NA           NA          1.2           15
Approach..........................          4.5           30          4.0           30          2.3           34
----------------------------------------------------------------------------------------------------------------

    (g) Engines comply with an applicable standard if the testing 
results show that the engine type certificate family's characteristic 
level does not exceed the numerical level of that standard, as described 
in the applicable appendix of Annex 16.
    (h) The system and procedure for sampling and measurement of gaseous 
emissions shall be as specified by in Appendices 2, 3, 4, 5 and 6 to the 
International Civil Aviation Organization (ICAO) Annex 16, Environmental 
Protection, Volume II, Aircraft Engine Emissions, Third Edition, July 
2008. This incorporation by reference was approved by the Director of 
the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 
51. This document can be obtained from the ICAO, Document Sales Unit, 
999 University Street, Montreal, Quebec H3C 5H7, Canada, phone +1 514-
954-8022, or www.icao.int or sales25icao.int. Copies can be reviewed at 
the FAA New England Regional Office, 12 New England Executive Park, 
Burlington, Massachusetts, 781-238-7101, or at the National Archives and 
Records Administration (NARA). For information on the availability of 
this material at NARA, call 202-741-6030, or go to: http://
www.archives.gov/federal--register/code--of--federal--regulations/ibr--
locations.html.

[Doc. No. FAA-2012-1333, 77 FR 76853, Dec. 31, 2012]

[[Page 927]]



Sec. Sec. 34.61-31.71  [Reserved]



PART 35_AIRWORTHINESS STANDARDS: PROPELLERS--Table of Contents



                            Subpart A_General

Sec.
35.1 Applicability.
35.2 Propeller configuration.
35.3 Instructions for propeller installation and operation.
35.4 Instructions for Continued Airworthiness.
35.5 Propeller ratings and operating limitations.
35.7 Features and characteristics.

                    Subpart B_Design and Construction

35.11 [Reserved]
35.13 [Reserved]
35.15 Safety analysis.
35.16 Propeller critical parts.
35.17 Materials and manufacturing methods.
35.19 Durability.
35.21 Variable and reversible pitch propellers.
35.22 Feathering propellers.
35.23 Propeller control system.
35.24 Strength.

                     Subpart C_Tests and Inspections

35.31 [Reserved]
35.33 General.
35.34 Inspections, adjustments and repairs.
35.35 Centrifugal load tests.
35.36 Bird impact.
35.37 Fatigue limits and evaluation.
35.38 Lightning strike.
35.39 Endurance test.
35.40 Functional test.
35.41 Overspeed and overtorque.
35.42 Components of the propeller control system.
35.43 Propeller hydraulic components.
35.45 [Reserved]
35.47 [Reserved]

Appendix A to Part 35--Instructions for Continued Airworthiness

    Authority: 49 U.S.C. 106(g), 40113, 44701-44702, 44704.

    Source: Docket No. 2095, 29 FR 7458, June 10, 1964, unless otherwise 
noted.



                            Subpart A_General



Sec. 35.1  Applicability.

    (a) This part prescribes airworthiness standards for the issue of 
type certificates and changes to those certificates, for propellers.
    (b) Each person who applies under part 21 for such a certificate or 
change must show compliance with the applicable requirements of this 
part.
    (c) An applicant is eligible for a propeller type certificate and 
changes to those certificates after demonstrating compliance with 
subparts A, B and C of this part. However, the propeller may not be 
installed on an airplane unless the applicant has shown compliance with 
either Sec. 23.907 or Sec. 25.907 of this chapter, as applicable, or 
compliance is not required for installation on that airplane.
    (d) For the purposes of this part, the propeller consists of those 
components listed in the propeller type design, and the propeller system 
consists of the propeller and all the components necessary for its 
functioning, but not necessarily included in the propeller type design.

[Amdt. 35-3, 41 FR 55475, Dec. 20, 1976, as amended by Amdt. 35-8, 73 FR 
63346, Oct. 24, 2008]



Sec. 35.2  Propeller configuration.

    The applicant must provide a list of all the components, including 
references to the relevant drawings and software design data, that 
define the type design of the propeller to be approved under Sec. 21.31 
of this chapter.

[Amdt. 35-8, 73 FR 63346, Oct. 24, 2008]



Sec. 35.3  Instructions for propeller installation and operation.

    The applicant must provide instructions that are approved by the 
Administrator. Those approved instructions must contain:
    (a) Instructions for installing the propeller, which:
    (1) Include a description of the operational modes of the propeller 
control system and functional interface of the control system with the 
airplane and engine systems;
    (2) Specify the physical and functional interfaces with the 
airplane, airplane equipment and engine;
    (3) Define the limiting conditions on the interfaces from paragraph 
(a)(2) of this section;
    (4) List the limitations established under Sec. 35.5;

[[Page 928]]

    (5) Define the hydraulic fluids approved for use with the propeller, 
including grade and specification, related operating pressure, and 
filtration levels; and
    (6) State the assumptions made to comply with the requirements of 
this part.
    (b) Instructions for operating the propeller which must specify all 
procedures necessary for operating the propeller within the limitations 
of the propeller type design.

[Amdt. 35-8, 73 FR 63346, Oct. 24, 2008]



Sec. 35.4  Instructions for Continued Airworthiness.

    The applicant must prepare Instructions for Continued Airworthiness 
in accordance with appendix A to this part that are acceptable to the 
Administrator. The instructions may be incomplete at type certification 
if a program exists to ensure their completion prior to delivery of the 
first aircraft with the propeller installed, or upon issuance of a 
standard certificate of airworthiness for an aircraft with the propeller 
installed, whichever occurs later.

[Amdt. 35-5, 45 FR 60181, Sept. 11, 1980]



Sec. 35.5  Propeller ratings and operating limitations.

    (a) Propeller ratings and operating limitations must:
    (1) Be established by the applicant and approved by the 
Administrator.
    (2) Be included directly or by reference in the propeller type 
certificate data sheet, as specified in Sec. 21.41 of this chapter.
    (3) Be based on the operating conditions demonstrated during the 
tests required by this part as well as any other information the 
Administrator requires as necessary for the safe operation of the 
propeller.
    (b) Propeller ratings and operating limitations must be established 
for the following, as applicable:
    (1) Power and rotational speed:
    (i) For takeoff.
    (ii) For maximum continuous.
    (iii) If requested by the applicant, other ratings may also be 
established.
    (2) Overspeed and overtorque limits.

[Amdt. 35-8, 73 FR 63346, Oct. 24, 2008]



Sec. 35.7  Features and characteristics.

    (a) The propeller may not have features or characteristics, revealed 
by any test or analysis or known to the applicant, that make it unsafe 
for the uses for which certification is requested.
    (b) If a failure occurs during a certification test, the applicant 
must determine the cause and assess the effect on the airworthiness of 
the propeller. The applicant must make changes to the design and conduct 
additional tests that the Administrator finds necessary to establish the 
airworthiness of the propeller.

[Amdt. 35-8, 73 FR 63346, Oct. 24, 2008]



                    Subpart B_Design and Construction



Sec. 35.11  [Reserved]



Sec. 35.13  [Reserved]



Sec. 35.15  Safety analysis.

    (a)(1) The applicant must analyze the propeller system to assess the 
likely consequences of all failures that can reasonably be expected to 
occur. This analysis will take into account, if applicable:
    (i) The propeller system in a typical installation. When the 
analysis depends on representative components, assumed interfaces, or 
assumed installed conditions, the assumptions must be stated in the 
analysis.
    (ii) Consequential secondary failures and dormant failures.
    (iii) Multiple failures referred to in paragraph (d) of this 
section, or that result in the hazardous propeller effects defined in 
paragraph (g)(1) of this section.
    (2) The applicant must summarize those failures that could result in 
major propeller effects or hazardous propeller effects defined in 
paragraph (g) of this section, and estimate the probability of 
occurrence of those effects.
    (3) The applicant must show that hazardous propeller effects are not 
predicted to occur at a rate in excess of that defined as extremely 
remote (probability of 10-7 or less per propeller

[[Page 929]]

flight hour). Since the estimated probability for individual failures 
may be insufficiently precise to enable the applicant to assess the 
total rate for hazardous propeller effects, compliance may be shown by 
demonstrating that the probability of a hazardous propeller effect 
arising from an individual failure can be predicted to be not greater 
than 10-8 per propeller flight hour. In dealing with 
probabilities of this low order of magnitude, absolute proof is not 
possible and reliance must be placed on engineering judgment and 
previous experience combined with sound design and test philosophies.
    (b) If significant doubt exists as to the effects of failures or 
likely combination of failures, the Administrator may require 
assumptions used in the analysis to be verified by test.
    (c) The primary failures of certain single propeller elements (for 
example, blades) cannot be sensibly estimated in numerical terms. If the 
failure of such elements is likely to result in hazardous propeller 
effects, those elements must be identified as propeller critical parts. 
For propeller critical parts, applicants must meet the prescribed 
integrity specifications of Sec. 35.16. These instances must be stated 
in the safety analysis.
    (d) If reliance is placed on a safety system to prevent a failure 
progressing to hazardous propeller effects, the possibility of a safety 
system failure in combination with a basic propeller failure must be 
included in the analysis. Such a safety system may include safety 
devices, instrumentation, early warning devices, maintenance checks, and 
other similar equipment or procedures. If items of the safety system are 
outside the control of the propeller manufacturer, the assumptions of 
the safety analysis with respect to the reliability of these parts must 
be clearly stated in the analysis and identified in the propeller 
installation and operation instructions required under Sec. 35.3.
    (e) If the safety analysis depends on one or more of the following 
items, those items must be identified in the analysis and appropriately 
substantiated.
    (1) Maintenance actions being carried out at stated intervals. This 
includes verifying that items that could fail in a latent manner are 
functioning properly. When necessary to prevent hazardous propeller 
effects, these maintenance actions and intervals must be published in 
the instructions for continued airworthiness required under Sec. 35.4. 
Additionally, if errors in maintenance of the propeller system could 
lead to hazardous propeller effects, the appropriate maintenance 
procedures must be included in the relevant propeller manuals.
    (2) Verification of the satisfactory functioning of safety or other 
devices at pre-flight or other stated periods. The details of this 
satisfactory functioning must be published in the appropriate manual.
    (3) The provision of specific instrumentation not otherwise 
required. Such instrumentation must be published in the appropriate 
documentation.
    (4) A fatigue assessment.
    (f) If applicable, the safety analysis must include, but not be 
limited to, assessment of indicating equipment, manual and automatic 
controls, governors and propeller control systems, synchrophasers, 
synchronizers, and propeller thrust reversal systems.
    (g) Unless otherwise approved by the Administrator and stated in the 
safety analysis, the following failure definitions apply to compliance 
with this part.
    (1) The following are regarded as hazardous propeller effects:
    (i) The development of excessive drag.
    (ii) A significant thrust in the opposite direction to that 
commanded by the pilot.
    (iii) The release of the propeller or any major portion of the 
propeller.
    (iv) A failure that results in excessive unbalance.
    (2) The following are regarded as major propeller effects for 
variable pitch propellers:
    (i) An inability to feather the propeller for feathering propellers.
    (ii) An inability to change propeller pitch when commanded.
    (iii) A significant uncommanded change in pitch.

[[Page 930]]

    (iv) A significant uncontrollable torque or speed fluctuation.

[Amdt. 35-8, 73 FR 63346, Oct. 24, 2008, as amended by Amdt. 35-9, 78 FR 
4041, Jan. 18, 2013; Amdt. 35-9A, 78 FR 45052, July 26, 2013]



Sec. 35.16  Propeller critical parts.

    The integrity of each propeller critical part identified by the 
safety analysis required by Sec. 35.15 must be established by:
    (a) A defined engineering process for ensuring the integrity of the 
propeller critical part throughout its service life,
    (b) A defined manufacturing process that identifies the requirements 
to consistently produce the propeller critical part as required by the 
engineering process, and
    (c) A defined service management process that identifies the 
continued airworthiness requirements of the propeller critical part as 
required by the engineering process.

[Amdt. 35-9, 78 FR 4042, Jan. 18, 2013]



Sec. 35.17  Materials and manufacturing methods.

    (a) The suitability and durability of materials used in the 
propeller must:
    (1) Be established on the basis of experience, tests, or both.
    (2) Account for environmental conditions expected in service.
    (b) All materials and manufacturing methods must conform to 
specifications acceptable to the Administrator.
    (c) The design values of properties of materials must be suitably 
related to the most adverse properties stated in the material 
specification for applicable conditions expected in service.

[Amdt. 35-8, 73 FR 63347, Oct. 24, 2008]



Sec. 35.19  Durability.

    Each part of the propeller must be designed and constructed to 
minimize the development of any unsafe condition of the propeller 
between overhaul periods.



Sec. 35.21  Variable and reversible pitch propellers.

    (a) No single failure or malfunction in the propeller system will 
result in unintended travel of the propeller blades to a position below 
the in-flight low-pitch position. The extent of any intended travel 
below the in-flight low-pitch position must be documented by the 
applicant in the appropriate manuals. Failure of structural elements 
need not be considered if the occurrence of such a failure is shown to 
be extremely remote under Sec. 35.15.
    (b) For propellers incorporating a method to select blade pitch 
below the in-flight low pitch position, provisions must be made to sense 
and indicate to the flight crew that the propeller blades are below that 
position by an amount defined in the installation manual. The method for 
sensing and indicating the propeller blade pitch position must be such 
that its failure does not affect the control of the propeller.

[Amdt. 35-8, 73 FR 63347, Oct. 24, 2008]



Sec. 35.22  Feathering propellers.

    (a) Feathering propellers are intended to feather from all flight 
conditions, taking into account expected wear and leakage. Any 
feathering and unfeathering limitations must be documented in the 
appropriate manuals.
    (b) Propeller pitch control systems that use engine oil to feather 
must incorporate a method to allow the propeller to feather if the 
engine oil system fails.
    (c) Feathering propellers must be designed to be capable of 
unfeathering after the propeller system has stabilized to the minimum 
declared outside air temperature.

[Amdt. 35-8, 73 FR 63347, Oct. 24, 2008]



Sec. 35.23  Propeller control system.

    The requirements of this section apply to any system or component 
that controls, limits or monitors propeller functions.
    (a) The propeller control system must be designed, constructed and 
validated to show that:
    (1) The propeller control system, operating in normal and 
alternative operating modes and in transition between operating modes, 
performs the functions defined by the applicant throughout the declared 
operating conditions and flight envelope.
    (2) The propeller control system functionality is not adversely 
affected

[[Page 931]]

by the declared environmental conditions, including temperature, 
electromagnetic interference (EMI), high intensity radiated fields 
(HIRF) and lightning. The environmental limits to which the system has 
been satisfactorily validated must be documented in the appropriate 
propeller manuals.
    (3) A method is provided to indicate that an operating mode change 
has occurred if flight crew action is required. In such an event, 
operating instructions must be provided in the appropriate manuals.
    (b) The propeller control system must be designed and constructed so 
that, in addition to compliance with Sec. 35.15:
    (1) No single failure or malfunction of electrical or electronic 
components in the control system results in a hazardous propeller 
effect.
    (2) Failures or malfunctions directly affecting the propeller 
control system in a typical airplane, such as structural failures of 
attachments to the control, fire, or overheat, do not lead to a 
hazardous propeller effect.
    (3) The loss of normal propeller pitch control does not cause a 
hazardous propeller effect under the intended operating conditions.
    (4) The failure or corruption of data or signals shared across 
propellers does not cause a hazardous propeller effect.
    (c) Electronic propeller control system imbedded software must be 
designed and implemented by a method approved by the Administrator that 
is consistent with the criticality of the performed functions and that 
minimizes the existence of software errors.
    (d) The propeller control system must be designed and constructed so 
that the failure or corruption of airplane-supplied data does not result 
in hazardous propeller effects.
    (e) The propeller control system must be designed and constructed so 
that the loss, interruption or abnormal characteristic of airplane-
supplied electrical power does not result in hazardous propeller 
effects. The power quality requirements must be described in the 
appropriate manuals.

[Amdt. 35-8, 73 FR 63347, Oct. 24, 2008]



Sec. 35.24  Strength.

    The maximum stresses developed in the propeller may not exceed 
values acceptable to the Administrator considering the particular form 
of construction and the most severe operating conditions.

[Amdt. 35-8, 73 FR 63348, Oct. 24, 2008]



                     Subpart C_Tests and Inspections



Sec. 35.31  [Reserved]



Sec. 35.33  General.

    (a) Each applicant must furnish test article(s) and suitable testing 
facilities, including equipment and competent personnel, and conduct the 
required tests in accordance with part 21 of this chapter.
    (b) All automatic controls and safety systems must be in operation 
unless it is accepted by the Administrator as impossible or not required 
because of the nature of the test. If needed for substantiation, the 
applicant may test a different propeller configuration if this does not 
constitute a less severe test.
    (c) Any systems or components that cannot be adequately 
substantiated by the applicant to the requirements of this part are 
required to undergo additional tests or analysis to demonstrate that the 
systems or components are able to perform their intended functions in 
all declared environmental and operating conditions.

[Amdt. 35-8, 73 FR 63348, Oct. 24, 2008]



Sec. 35.34  Inspections, adjustments and repairs.

    (a) Before and after conducting the tests prescribed in this part, 
the test article must be subjected to an inspection, and a record must 
be made of all the relevant parameters, calibrations and settings.
    (b) During all tests, only servicing and minor repairs are 
permitted. If major repairs or part replacement is required, the 
Administrator must approve the repair or part replacement

[[Page 932]]

prior to implementation and may require additional testing. Any 
unscheduled repair or action on the test article must be recorded and 
reported.

[Amdt. 35-8, 73 FR 63348, Oct. 24, 2008]



Sec. 35.35  Centrifugal load tests.

    The applicant must demonstrate that a propeller complies with 
paragraphs (a), (b) and (c) of this section without evidence of failure, 
malfunction, or permanent deformation that would result in a major or 
hazardous propeller effect. When the propeller could be sensitive to 
environmental degradation in service, this must be considered. This 
section does not apply to fixed-pitch wood or fixed-pitch metal 
propellers of conventional design.
    (a) The hub, blade retention system, and counterweights must be 
tested for a period of one hour to a load equivalent to twice the 
maximum centrifugal load to which the propeller would be subjected 
during operation at the maximum rated rotational speed.
    (b) Blade features associated with transitions to the retention 
system (for example, a composite blade bonded to a metallic retention) 
must be tested either during the test of paragraph (a) of this section 
or in a separate component test for a period of one hour to a load 
equivalent to twice the maximum centrifugal load to which the propeller 
would be subjected during operation at the maximum rated rotational 
speed.
    (c) Components used with or attached to the propeller (for example, 
spinners, de-icing equipment, and blade erosion shields) must be 
subjected to a load equivalent to 159 percent of the maximum centrifugal 
load to which the component would be subjected during operation at the 
maximum rated rotational speed. This must be performed by either:
    (1) Testing at the required load for a period of 30 minutes; or
    (2) Analysis based on test.

[Amdt. 35-8, 73 FR 63348, Oct. 24, 2008]



Sec. 35.36  Bird impact.

    The applicant must demonstrate, by tests or analysis based on tests 
or experience on similar designs, that the propeller can withstand the 
impact of a 4-pound bird at the critical location(s) and critical flight 
condition(s) of a typical installation without causing a major or 
hazardous propeller effect. This section does not apply to fixed-pitch 
wood propellers of conventional design.

[Amdt. 35-8, 73 FR 63348, Oct. 24, 2008]



Sec. 35.37  Fatigue limits and evaluation.

    This section does not apply to fixed-pitch wood propellers of 
conventional design.
    (a) Fatigue limits must be established by tests, or analysis based 
on tests, for propeller:
    (1) Hubs.
    (2) Blades.
    (3) Blade retention components.
    (4) Components which are affected by fatigue loads and which are 
shown under Sec. 35.15 to have a fatigue failure mode leading to 
hazardous propeller effects.
    (b) The fatigue limits must take into account:
    (1) All known and reasonably foreseeable vibration and cyclic load 
patterns that are expected in service; and
    (2) Expected service deterioration, variations in material 
properties, manufacturing variations, and environmental effects.
    (c) A fatigue evaluation of the propeller must be conducted to show 
that hazardous propeller effects due to fatigue will be avoided 
throughout the intended operational life of the propeller on either:
    (1) The intended airplane by complying with Sec. 23.907 or Sec. 
25.907 of this chapter, as applicable; or
    (2) A typical airplane.

[Amdt. 35-8, 73 FR 63348, Oct. 24, 2008]



Sec. 35.38  Lightning strike.

    The applicant must demonstrate, by tests, analysis based on tests, 
or experience on similar designs, that the propeller can withstand a 
lightning strike without causing a major or hazardous propeller effect. 
The limit to which the propeller has been qualified must be documented 
in the appropriate manuals. This section does not apply to fixed-pitch 
wood propellers of conventional design.

[Amdt. 35-8, 73 FR 63348, Oct. 24, 2008]

[[Page 933]]



Sec. 35.39  Endurance test.

    Endurance tests on the propeller system must be made on a 
representative engine in accordance with paragraph (a) or (b) of this 
section, as applicable, without evidence of failure or malfunction.
    (a) Fixed-pitch and ground adjustable-pitch propellers must be 
subjected to one of the following tests:
    (1) A 50-hour flight test in level flight or in climb. The propeller 
must be operated at takeoff power and rated rotational speed during at 
least five hours of this flight test, and at not less than 90 percent of 
the rated rotational speed for the remainder of the 50 hours.
    (2) A 50-hour ground test at takeoff power and rated rotational 
speed.
    (b) Variable-pitch propellers must be subjected to one of the 
following tests:
    (1) A 110-hour endurance test that must include the following 
conditions:
    (i) Five hours at takeoff power and rotational speed and thirty 10-
minute cycles composed of:
    (A) Acceleration from idle,
    (B) Five minutes at takeoff power and rotational speed,
    (C) Deceleration, and
    (D) Five minutes at idle.
    (ii) Fifty hours at maximum continuous power and rotational speed,
    (iii) Fifty hours, consisting of ten 5-hour cycles composed of:
    (A) Five accelerations and decelerations between idle and takeoff 
power and rotational speed,
    (B) Four and one half hours at approximately even incremental 
conditions from idle up to, but not including, maximum continuous power 
and rotational speed, and
    (C) Thirty minutes at idle.
    (2) The operation of the propeller throughout the engine endurance 
tests prescribed in part 33 of this chapter.
    (c) An analysis based on tests of propellers of similar design may 
be used in place of the tests of paragraphs (a) and (b) of this section.

[Amdt. 35-8, 73 FR 63348, Oct. 24, 2008]



Sec. 35.40  Functional test.

    The variable-pitch propeller system must be subjected to the 
applicable functional tests of this section. The same propeller system 
used in the endurance test (Sec. 35.39) must be used in the functional 
tests and must be driven by a representative engine on a test stand or 
on an airplane. The propeller must complete these tests without evidence 
of failure or malfunction. This test may be combined with the endurance 
test for accumulation of cycles.
    (a) Manually-controllable propellers. Five hundred representative 
flight cycles must be made across the range of pitch and rotational 
speed.
    (b) Governing propellers. Fifteen hundred complete cycles must be 
made across the range of pitch and rotational speed.
    (c) Feathering propellers. Fifty cycles of feather and unfeather 
operation must be made.
    (d) Reversible-pitch propellers. Two hundred complete cycles of 
control must be made from lowest normal pitch to maximum reverse pitch. 
During each cycle, the propeller must run for 30 seconds at the maximum 
power and rotational speed selected by the applicant for maximum reverse 
pitch.
    (e) An analysis based on tests of propellers of similar design may 
be used in place of the tests of this section.

[Amdt. 35-8, 73 FR 63349, Oct. 24, 2008]



Sec. 35.41  Overspeed and overtorque.

    (a) When the applicant seeks approval of a transient maximum 
propeller overspeed, the applicant must demonstrate that the propeller 
is capable of further operation without maintenance action at the 
maximum propeller overspeed condition. This may be accomplished by:
    (1) Performance of 20 runs, each of 30 seconds duration, at the 
maximum propeller overspeed condition; or
    (2) Analysis based on test or service experience.
    (b) When the applicant seeks approval of a transient maximum 
propeller overtorque, the applicant must demonstrate that the propeller 
is capable of further operation without maintenance action at the 
maximum propeller overtorque condition. This may be accomplished by:
    (1) Performance of 20 runs, each of 30 seconds duration, at the 
maximum propeller overtorque condition; or

[[Page 934]]

    (2) Analysis based on test or service experience.

[Amdt. 35-8, 73 FR 63349, Oct. 24, 2008]



Sec. 35.42  Components of the propeller control system.

    The applicant must demonstrate by tests, analysis based on tests, or 
service experience on similar components, that each propeller blade 
pitch control system component, including governors, pitch change 
assemblies, pitch locks, mechanical stops, and feathering system 
components, can withstand cyclic operation that simulates the normal 
load and pitch change travel to which the component would be subjected 
during the initially declared overhaul period or during a minimum of 
1,000 hours of typical operation in service.

[Amdt. 35-8, 73 FR 63349, Oct. 24, 2008]



Sec. 35.43  Propeller hydraulic components.

    Applicants must show by test, validated analysis, or both, that 
propeller components that contain hydraulic pressure and whose 
structural failure or leakage from a structural failure could cause a 
hazardous propeller effect demonstrate structural integrity by:
    (a) A proof pressure test to 1.5 times the maximum operating 
pressure for one minute without permanent deformation or leakage that 
would prevent performance of the intended function.
    (b) A burst pressure test to 2.0 times the maximum operating 
pressure for one minute without failure. Leakage is permitted and seals 
may be excluded from the test.

[Amdt. 35-8, 73 FR 63349, Oct. 24, 2008]



Sec. 35.45  [Reserved]



Sec. 35.47  [Reserved]



  Sec. Appendix A to Part 35--Instructions for Continued Airworthiness

                              a35.1 general

    (a) This appendix specifies requirements for the preparation of 
Instructions for Continued Airworthiness as required by Sec. 35.4.
    (b) The Instructions for Continued Airworthiness for each propeller 
must include the Instructions for Continued Airworthiness for all 
propeller parts. If Instructions for Continued Airworthiness are not 
supplied by the propeller part manufacturer for a propeller part, the 
Instructions for Continued Airworthiness for the propeller must include 
the information essential to the continued airworthiness of the 
propeller.
    (c) The applicant must submit to the FAA a program to show how 
changes to the Instructions for Continued Airworthiness made by the 
applicant or by the manufacturers of propeller parts will be 
distributed.

                              a35.2 format

    (a) The Instructions for Continued Airworthiness must be in the form 
of a manual or manuals as appropriate for the quantity of data to be 
provided.
    (b) The format of the manual or manuals must provide for a practical 
arrangement.

                              a35.3 content

    The contents of the manual must be prepared in the English language. 
The Instructions for Continued Airworthiness must contain the following 
sections and information:
    (a) Propeller Maintenance Section. (1) Introduction information that 
includes an explanation of the propeller's features and data to the 
extent necessary for maintenance or preventive maintenance.
    (2) A detailed description of the propeller and its systems and 
installations.
    (3) Basic control and operation information describing how the 
propeller components and systems are controlled and how they operate, 
including any special procedures that apply.
    (4) Instructions for uncrating, acceptance checking, lifting, and 
installing the propeller.
    (5) Instructions for propeller operational checks.
    (6) Scheduling information for each part of the propeller that 
provides the recommended periods at which it should be cleaned, 
adjusted, and tested, the applicable wear tolerances, and the degree of 
work recommended at these periods. However, the applicant may refer to 
an accessory, instrument, or equipment manufacturer as the source of 
this information if it shows that the item has an exceptionally high 
degree of complexity requiring specialized maintenance techniques, test 
equipment, or expertise. The recommended overhaul periods and necessary 
cross-references to the Airworthiness Limitations section of the manual 
must also be included. In addition, the applicant must include an 
inspection program that includes the frequency and extent of the 
inspections necessary to provide for the continued airworthiness of the 
propeller.
    (7) Troubleshooting information describing probable malfunctions, 
how to recognize

[[Page 935]]

those malfunctions, and the remedial action for those malfunctions.
    (8) Information describing the order and method of removing and 
replacing propeller parts with any necessary precautions to be taken.
    (9) A list of the special tools needed for maintenance other than 
for overhauls.
    (b) Propeller Overhaul Section. (1) Disassembly information 
including the order and method of disassembly for overhaul.
    (2) Cleaning and inspection instructions that cover the materials 
and apparatus to be used and methods and precautions to be taken during 
overhaul. Methods of overhaul inspection must also be included.
    (3) Details of all fits and clearances relevant to overhaul.
    (4) Details of repair methods for worn or otherwise substandard 
parts and components along with information necessary to determine when 
replacement is necessary.
    (5) The order and method of assembly at overhaul.
    (6) Instructions for testing after overhaul.
    (7) Instructions for storage preparation including any storage 
limits.
    (8) A list of tools needed for overhaul.

                 a35.4 airworthiness limitations section

    The Instructions for Continued Airworthiness must contain a section 
titled Airworthiness Limitations that is segregated and clearly 
distinguishable from the rest of the document. This section must set 
forth each mandatory replacement time, inspection interval, and related 
procedure required for type certification. This section must contain a 
legible statement in a prominent location that reads: ``The 
Airworthiness Limitations section is FAA approved and specifies 
maintenance required under Sec. Sec. 43.16 and 91.403 of the Federal 
Aviation Regulations unless an alternative program has been FAA 
approved.''

[Amdt. 35-5, 45 FR 60182, Sept. 11, 1980, as amended by Amdt. 35-6, 54 
FR 34330, Aug. 18, 1989]



PART 36_NOISE STANDARDS: AIRCRAFT TYPE AND AIRWORTHINESS CERTIFICATION
--Table of Contents



                            Subpart A_General

Sec.
36.1 Applicability and definitions.
36.2 Requirements as of date of application.
36.3 Compatibility with airworthiness requirements.
36.5 Limitation of part.
36.6 Incorporations by reference.
36.7 Acoustical change: Transport category large airplanes and jet 
          airplanes.
36.9 Acoustical change: Propeller-driven small airplanes and propeller-
          driven commuter category airplanes.
36.11 Acoustical change: Helicopters.
36.13 Acoustical change: Tiltrotor aircraft.

     Subpart B_Transport Category Large Airplanes and Jet Airplanes

36.101 Noise measurement and evaluation.
36.103 Noise limits.
36.105 Flight Manual Statement of Chapter 4 equivalency.

Subpart C [Reserved]

   Subpart D_Noise Limits for Supersonic Transport Category Airplanes

36.301 Noise limits: Concorde.

Subpart E [Reserved]

    Subpart F_Propeller Driven Small Airplanes and Propeller-Driven, 
                       Commuter Category Airplanes

36.501 Noise limits.

Subpart G [Reserved]

                          Subpart H_Helicopters

36.801 Noise measurement.
36.803 Noise evaluation and calculation.
36.805 Noise limits.

Subparts I-J [Reserved]

                          Subpart K_Tiltrotors

36.1101 Noise measurement and evaluation.
36.1103 Noise limits.

Subparts L-N [Reserved]

     Subpart O_Documentation, Operating Limitations and Information

36.1501 Procedures, noise levels and other information.
36.1581 Manuals, markings, and placards.
36.1583 Noncomplying agricultural and fire fighting airplanes.

Appendix A to Part 36--Aircraft Noise Measurement and Evaluation Under 
          Sec. 36.101
Appendix B to Part 36--Noise Levels for Transport Category and Jet 
          Airplanes Under Sec. 36.103
Appendixes C-E to Part 36 [Reserved]
Appendix F to Part 36--Flyover Noise Requirements for Propeller-Driven 
          Small Airplane and Propeller-Driven, Commuter Category 
          Airplane Certification Tests Prior to December 22, 1988

[[Page 936]]

Appendix G to Part 36--Takeoff Noise Requirements for Propeller-Driven 
          Small Airplane and Propeller-Driven, Commuter Category 
          Airplane Certification Tests on or After December 22, 1988
Appendix H to Part 36--Noise Requirements for Helicopters Under Subpart 
          H
Appendix I to Part 36 [Reserved]
Appendix J to Part 36--Alternative Noise Certification Procedure For 
          Helicopters Under Subpart H Having A Maximum Certificated 
          Takeoff Weight Of Not More Than 7,000 Pounds
Appendix K to Part 36--Noise Requirements for Tiltrotors Under Subpart K

    Authority: 42 U.S.C. 4321 et seq.; 49 U.S.C. 106(g), 40113, 44701-
44702, 44704, 44715; sec. 305, Pub. L. 96-193, 94 Stat. 50, 57; E.O. 
11514, 35 FR 4247, 3 CFR, 1966-1970 Comp., p. 902.

    Source: Docket No. 9337, 34 FR 18364, Nov. 18, 1969, unless 
otherwise noted.



                            Subpart A_General



Sec. 36.1  Applicability and definitions.

    (a) This part prescribes noise standards for the issue of the 
following certificates:
    (1) Type certificates, and changes to those certificates, and 
standard airworthiness certificates, for subsonic transport category 
large airplanes, and for subsonic jet airplanes regardless of category.
    (2) Type certificates and changes to those certificates, standard 
airworthiness certificates, and restricted category airworthiness 
certificates, for propeller-driven, small airplanes, and for propeller-
driven, commuter category airplanes except those airplanes that are 
designed for ``agricultural aircraft operations'' (as defined in Sec. 
137.3 of this chapter, as effective on January 1, 1966) or for 
dispersing fire fighting materials to which Sec. 36.1583 of this part 
does not apply.
    (3) A type certificate and changes to that certificate, and standard 
airworthiness certificates, for Concorde airplanes.
    (4) Type certificates, and changes to those certificates, for 
helicopters except those helicopters that are designated exclusively for 
``argicultural aircraft operations'' (as defined in Sec. 137.3 of this 
chapter, as effective on January 1, 1966), for dispensing fire fighting 
materials, or for carrying external loads (as defined in Sec. 133.1(b) 
of this chapter, as effective on December 20, 1976).
    (5) Type certificates, changes to those certificates, and standard 
airworthiness certificates, for tiltrotors.
    (b) Each person who applies under Part 21 of this chapter for a type 
of airworthiness certificate specified in this part must show compliance 
with the applicable requirements of this part, in addition to the 
applicable airworthiness requirements of this chapter.
    (c) Each person who applies under Part 21 of this chapter for 
approval of an acoustical change described in Sec. 21.93(b) of this 
chapter must show that the aircraft complies with the applicable 
provisions of Sec. Sec. 36.7, 36.9, 36.11 or 36.13 of this part in 
addition to the applicable airworthiness requirements of this chapter.
    (d) Each person who applies for the original issue of a standard 
airworthiness certificate for a transport category large airplane or for 
a jet airplane under Sec. 21.183 must, regardless of date of 
application, show compliance with the following provisions of this part 
(including appendix B):
    (1) The provisions of this part in effect on December 1, 1969, for 
subsonic airplanes that have not had any flight time before--
    (i) December 1, 1973, for airplanes with maximum weights greater 
than 75,000 pounds, except for airplanes that are powered by Pratt & 
Whitney Turbo Wasp JT3D series engines;
    (ii) December 31, 1974, for airplanes with maximum weights greater 
than 75,000 pounds and that are powered by Pratt & Whitney Turbo Wasp 
JT3D series engines; and
    (iii) December 31, 1974, for airplanes with maximum weights of 
75,000 pounds and less.
    (2) The provisions of this part in effect on October 13, 1977, 
including the stage 2 noise limits, for Concorde airplanes that have not 
had flight time before January 1, 1980.
    (e) Each person who applies for the original issue of a standard 
airworthiness certificate under Sec. 21.183, or for the original issue 
of a restricted category airworthiness certificate under Sec. 21.185, 
for propeller-driven, commuter category airplanes for a propeller driven

[[Page 937]]

small airplane that has not had any flight time before January 1, 1980, 
must show compliance with the applicable provisions of this part.
    (f) For the purpose of showing compliance with this part for 
transport category large airplanes and jet airplanes regardless of 
category, the following terms have the following meanings:
    (1) A ``Stage 1 noise level'' means a flyover, lateral or approach 
noise level greater than the Stage 2 noise limits prescribed in section 
B36.5(b) of appendix B of this part.
    (2) A ``Stage 1 airplane'' means an airplane that has not been shown 
under this part to comply with the flyover, lateral, and approach noise 
levels required for Stage 2 or Stage 3 airplanes.
    (3) A ``Stage 2 noise level'' means a noise level at or below the 
Stage 2 noise limits prescribed in section B36.5(b) of appendix B of 
this part but higher than the Stage 3 noise limits prescribed in section 
B36.5(c) of appendix B of this part.
    (4) A ``Stage 2 airplane'' means an airplane that has been shown 
under this part to comply with Stage 2 noise levels prescribed in 
section B36.5(b) of appendix B of this part (including use of the 
applicable tradeoff provisions specified in section B36.6) and that does 
not comply with the requirements for a Stage 3 airplane.
    (5) A ``Stage 3 noise level'' means a noise level at or below the 
Stage 3 noise limits prescribed in section B36.5(c) of appendix B of 
this part.
    (6) A ``Stage 3 airplane'' means an airplane that has been shown 
under this part to comply with Stage 3 noise levels prescribed in 
section B36.5(c) of appendix B of this part (including use of the 
applicable tradeoff provisions specified in section B36.6).
    (7) A ``subsonic airplane'' means an airplane for which the maximum 
operating limit speed, Mmo, does not exceed a Mach number of 
1.
    (8) A ``supersonic airplane'' means an airplane for which the 
maximum operating limit speed, Mmo, exceeds a Mach number of 
1.
    (9) A ``Stage 4 noise level'' means a noise level at or below the 
Stage 4 noise limit prescribed in section B36.5(d) of appendix B of this 
part.
    (10) A ``Stage 4 airplane'' means an airplane that has been shown 
under this part not to exceed the Stage 4 noise limit prescribed in 
section B36.5(d) of appendix B of this part.
    (11) A ``Chapter 4 noise level'' means a noise level at or below the 
maximum noise level prescribed in Chapter 4, Paragraph 4.4, Maximum 
Noise Levels, of the International Civil Aviation Organization (ICAO) 
Annex 16, Volume I, Amendment 7, effective March 21, 2002. [Incorporated 
by reference, see Sec. 36.6].
    (g) For the purpose of showing compliance with this part for 
transport category large airplanes and jet airplanes regardless of 
category, each airplane may not be identified as complying with more 
than one stage or configuration simultaneously.
    (h) For the purpose of showing compliance with this part, for 
helicopters in the primary, normal, transport, and restricted 
categories, the following terms have the specified meanings:
    (1) Stage 1 noise level means a takeoff, flyover, or approach noise 
level greater than the Stage 2 noise limits prescribed in section 
H36.305 of appendix H of this part, or a flyover noise level greater 
than the Stage 2 noise limits prescribed in section J36.305 of appendix 
J of this part.
    (2) Stage 1 helicopter means a helicopter that has not been shown 
under this part to comply with the takeoff, flyover, and approach noise 
levels required for Stage 2 helicopters as prescribed in section H36.305 
of appendix H of this part, or a helicopter that has not been shown 
under this part to comply with the flyover noise level required for 
Stage 2 helicopters as prescribed in section J36.305 of appendix J of 
this part.
    (3) Stage 2 noise level means a takeoff, flyover, or approach noise 
level at or below the Stage 2 noise limits prescribed in section H36.305 
of appendix H of this part, or a flyover noise level at or below the 
Stage 2 limit prescribed in section J36.305 of appendix J of this part.
    (4) Stage 2 helicopter means a helicopter that has been shown under 
this part to comply with Stage 2 noise limits (including applicable 
tradeoffs) prescribed in section H36.305 of appendix H of this part, or 
a helicopter that has

[[Page 938]]

been shown under this part to comply with the Stage 2 noise limit 
prescribed in section J36.305 of appendix J of this part.
    (5) A ``Stage 3 noise level'' means a takeoff, flyover, or approach 
noise level at or below the Stage 3 noise limits prescribed in section 
H36.305 of appendix H of this part, or a flyover noise level at or below 
the Stage 3 noise limit prescribed in section J36.305 of appendix J of 
this part.
    (6) A ``Stage 3 helicopter'' means a helicopter that has been shown 
under this part to comply with the Stage 3 noise limits (including 
applicable tradeoffs) prescribed in section H36.305 of appendix H of 
this part, or a helicopter that has been shown under this part to comply 
with the Stage 3 noise limit prescribed in section J36.305 of appendix J 
of this part.
    (7) Maximum normal operating RPM means the highest rotor speed 
corresponding to the airworthiness limit imposed by the manufacturer and 
approved by the FAA. Where a tolerance on the highest rotor speed is 
specified, the maximum normal operating rotor speed is the highest rotor 
speed for which that tolerance is given. If the rotor speed is 
automatically linked with flight condition, the maximum normal operating 
rotor speed corresponding with the reference flight condition must be 
used during the noise certification procedure. If rotor speed can be 
changed by pilot action, the highest normal operating rotor speed 
specified in the flight manual limitation section for reference 
conditions must be used during the noise certification procedure.
    (i) For the purpose of showing compliance with this part for 
tiltrotors, the following terms have the specified meanings:
    Airplane mode means a configuration with nacelles on the down stops 
(axis aligned horizontally) and rotor speed set to cruise revolutions 
per minute (RPM).
    Airplane mode RPM means the lower range of rotor rotational speed in 
RPM defined for the airplane mode cruise flight condition.
    Fixed operation points mean designated nacelle angle positions 
selected for airworthiness reference. These are default positions used 
to refer to normal nacelle positioning operation of the aircraft. The 
nacelle angle is controlled by a self-centering switch. When the nacelle 
angle is 0 degrees (airplane mode) and the pilot moves the nacelle 
switch upwards, the nacelles are programmed to automatically turn to the 
first default position (for example, 60 degrees) where they will stop. A 
second upward move of the switch will tilt the nacelle to the second 
default position (for example, 75 degrees). Above the last default 
position, the nacelle angle can be set to any angle up to approximately 
95 degrees by moving the switch in the up or down direction. The number 
and position of the fixed operation points may vary on different 
tiltrotor configurations.
    Nacelle angle is defined as the angle between the rotor shaft 
centerline and the longitudinal axis of the aircraft fuselage.
    Tiltrotor means a class of aircraft capable of vertical take-off and 
landing, within the powered-lift category, with rotors mounted at or 
near the wing tips that vary in pitch from near vertical to near 
horizontal configuration relative to the wing and fuselage.
    Vertical takeoff and landing (VTOL) mode means the aircraft state or 
configuration having the rotors orientated with the axis of rotation in 
a vertical manner (i.e., nacelle angle of approximately 90 degrees) for 
vertical takeoff and landing operations.
    VCON is defined as the maximum authorized speed for any 
nacelle angle in VTOL/Conversion mode.
    VTOL/Conversion mode is all approved nacelle positions where the 
design operating rotor speed is used for hover operations.
    VTOL mode RPM means highest range of RPM that occur for takeoff, 
approach, hover, and conversion conditions.

[Doc. No. 13243, Amdt. 36-4, 40 FR 1034, Jan. 6, 1975]

    Editorial Note: For Federal Register citations affecting Sec. 36.1, 
see the List of CFR Sections Affected, which appears in the Finding Aids 
section of the printed volume and at www.fdsys.gov.

[[Page 939]]



Sec. 36.2  Requirements as of date of application.

    (a) Section 21.17 of this chapter notwithstanding, each person who 
applies for a type certificate for an aircraft covered by this part, 
must show that the aircraft meets the applicable requirements of this 
part that are effective on the date of application for that type 
certificate. When the time interval between the date of application for 
the type certificate and the issuance of the type certificate exceeds 5 
years, the applicant must show that the aircraft meets the applicable 
requirements of this part that were effective on a date, to be selected 
by the applicant, not earlier than 5 years before the issue of the type 
certificate.
    (b) Section 21.101(a) of this chapter notwithstanding, each person 
who applies for an acoustical change to a type design specified in Sec. 
21.93(b) of this chapter must show compliance with the applicable 
requirements of this part that are effective on the date of application 
for the change in type design. When the time interval between the date 
of application for the change in type design and the issuance of the 
amended or supplemental type certificate exceeds 5 years, the applicant 
must show that the aircraft meets the applicable requirements of this 
part that were effective on a date, to be selected by the applicant, not 
earlier than 5 years before the issue of the amended or supplemental 
type certificate.
    (c) If an applicant elects to comply with a standard in this part 
that was effective after the filing of the application for a type 
certificate or change to a type design, the election:
    (1) Must be approved by the FAA;
    (2) Must include standards adopted between the date of application 
and the date of the election;
    (3) May include other standards adopted after the standard elected 
by the applicant as determined by the FAA.

[Amdt. 36-54, 67 FR 45211, July 8, 2002; Amdt. 36-24, 67 FR 63195, Oct. 
10, 2002]



Sec. 36.3  Compatibility with airworthiness requirements.

    It must be shown that the aircraft meets the airworthiness 
regulations constituting the type certification basis of the aircraft 
under all conditions in which compliance with this part is shown, and 
that all procedures used in complying with this part, and all procedures 
and information for the flight crew developed under this part, are 
consistent with the airworthiness regulations constituting the type 
certification basis of the aircraft.

[Doc. No. 9337, 34 FR 18364, Nov. 18, 1969, as amended by Amdt. 36-14, 
53 FR 3540, Feb. 5, 1988]



Sec. 36.5  Limitation of part.

    Pursuant to 49 U.S.C. 1431(b)(4), the noise levels in this part have 
been determined to be as low as is economically reasonable, 
technologically practicable, and appropriate to the type of aircraft to 
which they apply. No determination is made, under this part, that these 
noise levels are or should be acceptable or unacceptable for operation 
at, into, or out of, any airport.



Sec. 36.6  Incorporation by reference.

    (a) General. This part prescribes certain standards and procedures 
which are not set forth in full text in the rule. Those standards and 
procedures are contained in published material which is reasonably 
available to the class of persons affected and has been approved for 
incorporation by reference by the Director of the Federal Register under 
5 U.S.C. 552 (a) and 1 CFR Part 51.
    (b) Incorporated matter. (1) Each publication, or part of a 
publication, which is referenced but not set forth in full-text in this 
part and which is identified in paragraph (c) of this section is hereby 
incorporated by reference and made a part of part 36 of this chapter 
with the approval of the Director of the Federal Register.
    (2) Incorporated matter which is subject to subsequent change is 
incorporated by reference according to the specific reference and to the 
identification statement. Adoption of any subsequent change in 
incorporated matter is made under Part 11 of this chapter and 1 CFR Part 
51.

[[Page 940]]

    (c) Identification statement. The complete title or description 
which identifies each published matter incorporated by reference in this 
part is as follows:
    (1) International Electrotechnical Commission (IEC) Publications. 
(i) IEC Publication No. 179, entitled ``Precision Sound Level Meters,'' 
dated 1973.
    (ii) IEC Publication No. 225, entitled ``Octave, Half-Octave, Third 
Octave Band Filters Intended for the Analysis of Sounds and 
Vibrations,'' dated 1966.
    (iii) IEC Publication No. 651, entitled ``Sound Level Meters,'' 
first edition, dated 1979.
    (iv) IEC Publication No. 561, entitled ``Electro-acoustical 
Measuring Equipment for Aircraft Noise Certification,'' first edition, 
dated 1976.
    (v) IEC Publication No. 804, entitled ``Integrating-averaging Sound 
Level Meters,'' first edition, dated 1985.
    (vi) IEC Publication 61094-3, entitled ``Measurement Microphones--
Part 3: Primary Method for Free-Field Calibration of Laboratory Standard 
Microphones by the Reciprocity Technique'', edition 1.0, dated 1995.
    (vii) IEC Publication 61094-4, entitled ``Measurement Microphones--
Part 4: Specifications for Working Standard Microphones'', edition 1.0, 
dated 1995.
    (viii) IEC Publication 61260, entitled ``Electroacoustics-Octave-
Band and Fractional-Octave-Band filters'', edition 1.0, dated 1995.
    (ix) IEC Publication 61265, entitled ``Instruments for Measurement 
of Aircraft Noise-Performance Requirements for Systems to Measure One-
Third-Octave-Band Sound pressure Levels in Noise Certification of 
Transport-Category Aeroplanes,'' edition 1.0, dated 1995.
    (x) IEC Publication 60942, entitled ``Electroacoustics--Sound 
Calibrators,'' edition 2.0, dated 1997.
    (2) Society of Automotive Engineers (SAE) Publications. (i) SAE ARP 
866A, entitled ``Standard Values at Atmospheric Absorption as a Function 
of Temperature and Humidity for Use in Evaluating Aircraft Flyover 
Noise,'' dated March 15, 1975.
    (3) International Standards and Recommended Practices entitled 
``Environmental Protection, Annex 16 to the Convention on International 
Civil Aviation, Volume I, Aircraft Noise'', Third Edition, July 1993, 
Amendment 7, effective March 21, 2002.
    (d) Availability for purchase. Published material incorporated by 
reference in this part may be purchased at the price established by the 
publisher or distributor at the following mailing addresses:
    (1) IEC publications.
    (i) International Electrotechnical Commission, 3, rue de Varembe, 
Case postale 131, 1211 Geneva 20, Switzerland.
    (ii) American National Standard Institute, 11 West 42nd Street, New 
York City, New York 10036.
    (2) SAE publications. Society of Automotive Engineers, Inc., 400 
Commonwealth Drive, Warrentown, Pennsylvania 15096.
    (3) ICAO publications. International Civil Aviation Organization 
(ICAO), Document Sales Unit, 999 University Street, Montreal, Quebec H3C 
5H7, Canada.
    (e) Availability for inspection. A copy of each publication 
incorporated by reference in this part is available for public 
inspection at the following locations:
    (1) U.S. Department of Transportation, Docket Operations, West 
Building Ground Floor, Room W12-140, 1200 New Jersey Avenue, SE., 
Washington, DC 20590.
    (2) The National Archives and Records Administration (NARA). For 
information on the availability of this material at NARA, call 202-741-
6030, or go to: http://www.archives.gov/federal--register/code--of--
federal--regulations/ibr--locations.html.
    (3) The respective Region Headquarters of the Federal Aviation 
Administration as follows:
    (i) New England Region Headquarters, 12 New England Executive Park, 
Burlington, Massachusetts 01803.
    (ii) Eastern Region Headquarters, Federal Building, John F. Kennedy 
(JFK) International Airport, Jamaica, New York 11430.
    (iii) Southern Region Headquarters, 1701 Columbia Avenue, College 
Park, Georgia, 30337.
    (iv) Great Lakes Region Headquarters, O'Hare Lake Office Center,

[[Page 941]]

2300 East Devon Avenue, Des Plaines, Illinois 60018.
    (v) Central Region Headquarters, Federal Building, 601 East 12th 
Street, Kanasa City Missouri 64106.
    (vi) Southwest Region Headquarters, 2601 Meacham Boulevard, Fort 
Worth, Texas, 76137-4298.
    (vii) Northwest Mountain Region Headquarters, 1601 Lind Avenue, 
Southwest, Renton, Washington 98055.
    (viii) Western-Pacific Region Headquarters, 15000 Aviation 
Boulevard, Hawthorne, California 92007.
    (ix) Alaskan Region Headquarters, 222 West 7th Avenue, 14, 
Anchorage, Alaska, 99513.
    (x) European Office Headquarters, 15, Rue de la Loi (3rd Floor), B-
1040 Brussels, Belgium.

[Amdt. 36-9, 43 FR 8739, Mar. 3, 1978, as amended by Amdt. 36-16, 53 FR 
47400, Nov. 22, 1988; Amdt. 36-20, 57 FR 42854, Sept. 16, 1992; Amdt. 
36-54, 67 FR 45212, July 8, 2002; Amdt. 36-24, 28 FR 1512, Jan. 10, 
2003; 68 FR 2402, Jan. 16, 2003; Amdt. 36-26, 70 FR 38748, July 5, 2005; 
72 FR 68475, Dec. 5, 2007]



Sec. 36.7  Acoustical change: Transport category large airplanes
and jet airplanes.

    (a) Applicability. This section applies to all transport category 
large airplanes and jet airplanes for which an acoustical change 
approval is applied for under Sec. 21.93(b) of this chapter.
    (b) General requirements. Except as otherwise specifically provided, 
for each airplane covered by this section, the acoustical change 
approval requirements are as follows:
    (1) In showing compliance, noise levels must be measured and 
evaluated in accordance with the applicable procedures and conditions 
prescribed in Appendix A of this part.
    (2) Compliance with the noise limits prescribed in section B36.5 of 
appendix B must be shown in accordance with the applicable provisions of 
sections B36.7 and B36.8 of appendix B of this part.
    (c) Stage 1 airplanes. For each Stage 1 airplane prior to the change 
in type design, in addition to the provisions of paragraph (b) of this 
section, the following apply:
    (1) If an airplane is a Stage 1 airplane prior to the change in type 
design, it may not, after the change in type design, exceed the noise 
levels created prior to the change in type design. The tradeoff 
provisions of section B36.6 of appendix B of this part may not be used 
to increase the Stage 1 noise levels, unless the aircraft qualifies as a 
Stage 2 airplane.
    (2) In addition, for an airplane for which application is made after 
September 17, 1971--
    (i) There may be no reduction in power or thrust below the highest 
airworthiness approved power or thrust, during the tests conducted 
before and after the change in type design; and
    (ii) During the flyover and lateral noise tests conducted before the 
change in type design, the quietest airworthiness approved configuration 
available for the highest approved takeoff weight must be used.
    (d) Stage 2 airplanes. If an airplane is a Stage 2 airplane prior to 
the change in type design, the following apply, in addition to the 
provisions of paragraph (b) of this section:
    (1) Airplanes with high bypass ratio jet engines. For an airplane 
that has jet engines with a bypass ratio of 2 or more before a change in 
type design--
    (i) The airplane, after the change in type design, may not exceed 
either (A) each Stage 3 noise limit by more than 3 EPNdB, or (B) each 
Stage 2 noise limit, whichever is lower:
    (ii) The tradeoff provisions of section B36.6 of appendix B of this 
part may be used in determining compliance under this paragraph with 
respect to the Stage 2 noise limit or to the Stage 3 plus 3 EPNdB noise 
limits, as applicable; and
    (iii) During the flyover and lateral noise test conducted before the 
change in type design, the quietest airworthiness approved configuration 
available for the highest approved takeoff weight must be used.
    (2) Airplanes that do not have high bypass ratio jet engines. For an 
airplane that does not have jet engines with a bypass ratio of 2 or more 
before a change in type design--
    (i) The airplane may not be a Stage 1 airplane after the change in 
type design; and
    (ii) During the flyover and lateral noise tests conducted before the 
change

[[Page 942]]

in type design, the quietest airworthiness approved configuration 
available for the highest approved takeoff weight must be used.
    (e) Stage 3 airplanes. If an airplane is a Stage 3 airplane prior to 
the change in type design, the following apply, in addition to the 
provisions of paragraph (b) of this section:
    (1) If compliance with Stage 3 noise levels is not required before 
the change in type design, the airplane must--
    (i) Be a Stage 2 airplane after the change in type design and 
compliance must be shown under the provisions of paragraph (d)(1) or 
(d)(2) of this section, as appropriate; or
    (ii) Remain a Stage 3 airplane after the change in type design. 
Compliance must be shown under the provisions of paragraph (e)(2) of 
this section.
    (2) If compliance with Stage 3 noise levels is required before the 
change in type design, the airplane must be a Stage 3 airplane after the 
change in type design.
    (3) Applications on or after [August 14, 1989.] The airplane must 
remain a Stage 3 airplane after the change in type design.
    (4) If an airplane is a Stage 3 airplane prior to a change in type 
design, and becomes a Stage 4 after the change in type design, the 
airplane must remain a Stage 4 airplane.
    (f) Stage 4 airplanes. If an airplane is a Stage 4 airplane prior to 
a change in type design, the airplane must remain a Stage 4 airplane 
after the change in type design.

[Amdt. 36-7, 42 FR 12371, Mar. 3, 1977; Amdt. 36-8, 43 FR 8730, Mar. 2, 
1978; Amdt. 36-10, 43 FR 28420, June 29, 1978; Amdt. 36-12, 46 FR 33464, 
June 29, 1981; Amdt. 36-15, 53 FR 16366, May 6, 1988; 53 FR 18950, May 
25, 1988; Amdt. 36-17, 54 FR 21042, May 15, 1989; Amdt. 36-54, 67 FR 
45212, July 8, 2002; Amdt. 36-26, 70 FR 38749, July 5, 2005]



Sec. 36.9  Acoustical change: Propeller-driven small airplanes
and propeller-driven commuter category airplanes.

    For propeller-driven small airplanes in the primary, normal, 
utility, acrobatic, transport, and restricted categories and for 
propeller-driven, commuter category airplanes for which an acoustical 
change approval is applied for under Sec. 21.93(b) of this chapter 
after January 1, 1975, the following apply:
    (a) If the airplane was type certificated under this part prior to a 
change in type design, it may not subsequently exceed the noise limits 
specified in Sec. 36.501 of this part.
    (b) If the airplane was not type certificated under this part prior 
to a change in type design, it may not exceed the higher of the two 
following values:
    (1) The noise limit specified in Sec. 36.501 of this part, or
    (2) The noise level created prior to the change in type design, 
measured and corrected as prescribed in Sec. 36.501 of this part.

[Amdt. 36-16, 53 FR 47400, Nov. 22, 1988; 53 FR 50157, Dec. 13, 1988; 
Amdt. 36-19, 57 FR 41369, Sept. 9, 1992]



Sec. 36.11  Acoustical change: Helicopters.

    This section applies to all helicopters in the primary, normal, 
transport, and restricted categories for which an acoustical change 
approval is applied for under Sec. 21.93(b) of this chapter on or after 
March 6, 1986. Compliance with the requirements of this section must be 
demonstrated under appendix H of this part, or, for helicopters having a 
maximum certificated takeoff weight of not more than 7,000 pounds, 
compliance with this section may be demonstrated under appendix J of 
this part.
    (a) General requirements. Except as otherwise provided, for 
helicopters covered by this section, the acoustical change approval 
requirements are as follows:
    (1) In showing compliance with the requirements of appendix H of 
this part, noise levels must be measured, evaluated, and calculated in 
accordance with the applicable procedures and conditions prescribed in 
parts B and C of appendix H of this part. For helicopters having a 
maximum certificated takeoff weight of not more than 7,000 pounds that 
alternatively demonstrate compliance under appendix J of this part, the 
flyover noise level prescribed in appendix J of this part must be 
measured, evaluated, and calculated in accordance with the applicable 
procedures and conditions prescribed in

[[Page 943]]

parts B and C of appendix J of this part.
    (2) Compliance with the noise limits prescribed in section H36.305 
of appendix H of this part must be shown in accordance with the 
applicable provisions of part D of appendix H of this part. For those 
helicopters that demonstrate compliance with the requirements of 
appendix J of this part, compliance with the noise levels prescribed in 
section J36.305 of appendix J of this part must be shown in accordance 
with the applicable provisions of part D of appendix J of this part.
    (b) Stage 1 helicopters. Except as provided in Sec. 36.805(c), for 
each Stage 1 helicopter prior to a change in type design, the helicopter 
noise levels may not, after a change in type design, exceed the noise 
levels specified in section H36.305(a)(1) of appendix H of this part 
where the demonstration of compliance is under appendix H of this part. 
The tradeoff provisions under section H36.305(b) of appendix H of this 
part may not be used to increase any Stage 1 noise level beyond these 
limits. If an applicant chooses to demonstrate compliance under appendix 
J of this part, for each Stage 1 helicopter prior to a change in type 
design, the helicopter noise levels may not, after a change in type 
design, exceed the Stage 2 noise levels specified in section J36.305(a) 
of appendix J of this part.
    (c) Stage 2 helicopters. For each helicopter that is Stage 2 prior 
to a change in type design, after a change in type design the helicopter 
must either:
    (1) Remain a Stage 2 helicopter; or
    (2) Comply with Stage 3 requirements and remain a Stage 3 helicopter 
thereafter.
    (d) Stage 3 helicopters. For a helicopter that is a Stage 3 
helicopter prior to a change in type design, the helicopter must remain 
a Stage 3 helicopter after a change in type design.

[Doc. No. 26910, 57 FR 42854, Sept. 16, 1992, as amended by Amdt. 36-25, 
69 FR 31234, June 2, 2004; Amdt. 36-30, 79 FR 12044, Mar. 4, 2014]



Sec. 36.13  Acoustical change: Tiltrotor aircraft.

    The following requirements apply to tiltrotors in any category for 
which an acoustical change approval is applied for under Sec. 21.93(b) 
of this chapter on or after March 11, 2013:
    (a) In showing compliance with Appendix K of this part, noise levels 
must be measured, evaluated, and calculated in accordance with the 
applicable procedures and conditions prescribed in Appendix K of this 
part.
    (b) Compliance with the noise limits prescribed in section K4 (Noise 
Limits) of Appendix K of this part must be shown in accordance with the 
applicable provisions of sections K2 (Noise Evaluation Measure), K3 
(Noise Measurement Reference Points), K6 (Noise Certification Reference 
Procedures), and K7 (Test Procedures) of Appendix K of this part.
    (c) After a change in type design, tiltrotor noise levels may not 
exceed the limits specified in Sec. 36.1103.

[Amdt. 36-29, 78 FR 1139, Jan. 8, 2013]



     Subpart B_Transport Category Large Airplanes and Jet Airplanes



Sec. 36.101  Noise measurement and evaluation.

    For transport category large airplanes and jet airplanes, the noise 
generated by the airplane must be measured and evaluated under appendix 
A of this part or under an approved equivalent procedure.

[Amdt. 36-54, 67 FR 45212, July 8, 2002]



Sec. 36.103  Noise limits.

    (a) For subsonic transport category large airplanes and subsonic jet 
airplanes compliance with this section must be shown with noise levels 
measured and evaluated as prescribed in appendix A of this part, and 
demonstrated at the measuring points, and in accordance with the test 
procedures under section B36.8 (or an approved equivalent procedure), 
stated under appendix B of this part.
    (b) Type certification applications between November 5, 1975 and 
December 31, 2005. If application is made on or after November 5, 1975, 
and before January 1, 2006, it must be shown that the noise levels of 
the airplane are no greater than the Stage 3 noise limit prescribed in 
section B36.5(c) of appendix B of this part.

[[Page 944]]

    (c) Type certification applications on or after January 1, 2006. If 
application is made on or after January 1, 2006, it must be shown that 
the noise levels of the airplane are no greater than the Stage 4 noise 
limit prescribed in section B36.5(d) of appendix B of this part. Prior 
to January 1, 2006, an applicant may seek voluntary certification to 
Stage 4. If Stage 4 certification is chosen, the requirements of Sec. 
36.7(f) of this part will apply.

[Amdt. 36-54, 67 FR 45212, July 8, 2002, as amended by Amdt. 36-26, 70 
FR 38749, July 5, 2005]



Sec. 36.105  Flight Manual Statement of Chapter 4 equivalency.

    For each airplane that meets the requirements for Stage 4 
certification, the Airplane Flight Manual or operations manual must 
include the following statement: ``The following noise levels comply 
with part 36, Appendix B, Stage 4 maximum noise level requirements and 
were obtained by analysis of approved data from noise tests conducted 
under the provisions of part 36, Amendment 36 (insert part 36 amendment 
to which the airplane was certificated). The noise measurement and 
evaluation procedures used to obtain these noise levels are considered 
by the FAA to be equivalent to the Chapter 4 noise level required by the 
International Civil Aviation Organization (ICAO) in Annex 16, Volume I, 
Appendix 2, Amendment 7, effective March 21, 2002.'' [Incorporated by 
reference, see Sec. 36.6].

[Amdt. 36-26, 70 FR 38749, July 5, 2005; 70 FR 41610, July 20, 2005]

Subpart C [Reserved]



   Subpart D_Noise Limits for Supersonic Transport Category Airplanes



Sec. 36.301  Noise limits: Concorde.

    (a) General. For the Concorde airplane, compliance with this subpart 
must be shown with noise levels measured and evaluated as prescribed in 
Subpart B of this part, and demonstrated at the measuring points 
prescribed in appendix B of this part.
    (b) Noise limits. It must be shown, in accordance with the 
provisions of this part in effect on October 13, 1977, that the noise 
levels of the airplane are reduced to the lowest levels that are 
economically reasonable, technologically practicable, and appropriate 
for the Concorde type design.

[Amdt. 36-10, 43 FR 28420, June 29, 1978, as amended by Amdt. 36-54, 67 
FR 45212, July 8, 2002]

Subpart E [Reserved]



    Subpart F_Propeller Driven Small Airplanes and Propeller-Driven, 
                       Commuter Category Airplanes



Sec. 36.501  Noise limits.

    (a) Compliance with this subpart must be shown for--
    (1) Propeller driven small airplanes for which application for the 
issuance of a new, amended, or supplemental type certificate in the 
normal, utility, acrobatic, transport, or restricted category is made on 
or after October 10, 1973; and propeller-driven, commuter category 
airplanes for which application for the issuance of a type certificate 
in the commuter category is made on or after January 15, 1987.
    (2) Propeller driven small airplanes and propeller-driven, commuter 
category airplanes for which application is made for the original 
issuance of a standard airworthiness certificate or restricted category 
airworthiness certificate, and that have not had any flight time before 
January 1, 1980 (regardless of date of application).
    (3) Airplanes in the primary category:
    (i) Except as provided in paragraph (a)(3)(ii) of this section, for 
an airplane for which application for a type certificate in the primary 
category is made, and that was not previously certificated under 
appendix F of this part, compliance with appendix G of this part must be 
shown.
    (ii) For an airplane in the normal, utility or acrobatic category 
that (A) has a type certificate issued under this

[[Page 945]]

chapter, (B) has a standard airworthiness certificate issued under this 
chapter, (C) has not undergone an acoustical change from its type 
design, (D) has not previously been certificated under appendix F or G 
of this part, and (E) for which application for conversion to the 
primary category is made, no further showing of compliance with this 
part is required.
    (b) For aircraft covered by this subpart for which certification 
tests are completed before December 22, 1988, compliance must be shown 
with noise levels as measured and prescribed in Parts B and C of 
appendix F, or under approved equivalent procedures. It must be shown 
that the noise level of the airplane is no greater than the applicable 
limit set in Part D of appendix F.
    (c) For aircraft covered by this subpart for which certification 
tests are not completed before December 22, 1988, compliance must be 
shown with noise levels as measured and prescribed in Parts B and C of 
appendix G, or under approved equivalent procedures. It must be shown 
that the noise level of the airplane is no greater than the applicable 
limits set in Part D of appendix G.

[Doc. No. 13243, 40 FR 1034, Jan. 6, 1975, as amended by Amdt. 36-13, 52 
FR 1836, Jan. 15, 1987; Amdt. 36-16, 53 FR 47400, Nov. 22, 1988; Amdt. 
36-19, 57 FR 41369, Sept. 9, 1992]

Subpart G [Reserved]



                          Subpart H_Helicopters

    Source: Amdt. 36-14, 53 FR 3540, Feb. 5, 1988; 53 FR 7728, Mar. 10, 
1988, unless otherwise noted.



Sec. 36.801  Noise measurement.

    For primary, normal, transport, or restricted category helicopters 
for which certification is sought under appendix H of this part, the 
noise generated by the helicopter must be measured at the noise 
measuring points and under the test conditions prescribed in part B of 
appendix H of this part, or under an FAA-approved equivalent procedure. 
For those primary, normal, transport, and restricted category 
helicopters having a maximum certificated takeoff weight of not more 
than 7,000 pounds for which compliance with appendix J of this part is 
demonstrated, the noise generated by the helicopter must be measured at 
the noise measuring point and under the test conditions prescribed in 
part B of appendix J of this part, or an FAA-approved equivalent 
procedure.

[Doc. No. 26910, 57 FR 42854, Sept. 16, 1992, as amended by Amdt. 36-25, 
69 FR 31234, June 2, 2004]



Sec. 36.803  Noise evaluation and calculation.

    The noise measurement data required under Sec. 36.801 and obtained 
under appendix H of this part must be corrected to the reference 
conditions contained in part A of appendix H of this part, and evaluated 
under the procedures of part C of appendix H of this part, or an FAA-
approved equivalent procedure. The noise measurement data required under 
Sec. 36.801 and obtained under appendix J of this part must be 
corrected to the reference conditions contained in part A of appendix J 
of this part, and evaluated under the procedures of part C of appendix J 
of this part, or an FAA-approved equivalent procedure.

[Doc. No. 26910, 57 FR 42854, Sept. 16, 1992]



Sec. 36.805  Noise limits.

    (a) Compliance with the noise levels prescribed under part D of 
appendix H of this part, or under part D of appendix J of this part, 
must be shown for helicopters for which application for issuance of a 
type certificate in the primary, normal, transport, or restricted 
category is made on or after March 6, 1986.
    (b) For helicopters covered by this section, except as provided in 
paragraph (c) or (d)(2) of this section, it must be shown either:
    (1) When an application for issuance of a type certificate in the 
primary, normal, transport, or restricted category is made on and after 
March 6, 1986 and before May 5, 2014, that the noise levels of the 
helicopter are no greater than the Stage 2 noise limits prescribed in 
either section H36.305 of appendix H of this part or section J36.305 of 
appendix J of this part, as applicable; or

[[Page 946]]

    (2) When an application for issuance of a type certificate in the 
primary, normal, transport, or restricted category is made on or after 
May 5, 2014, that the noise levels of the helicopter are no greater than 
the Stage 3 noise limits prescribed in either section H36.305 of 
appendix H of this part, or section J36.305 of appendix J of this part, 
as applicable.
    (c) For helicopters for which application for issuance of an 
original type certificate in the primary, normal, transport, or 
restricted category is made on or after March 6, 1986, and which the FAA 
finds to be the first civil version of a helicopter that was designed 
and constructed for, and accepted for operational use by, an Armed Force 
of the United States or the U.S. Coast Guard on or before March 6, 1986, 
it must be shown that the noise levels of the helicopter are no greater 
than the noise limits for a change in type design as specified in 
section H36.305(a)(1)(ii) of appendix H of this part for compliance 
demonstrated under appendix H of this part, or as specified in section 
J36.305 of appendix J of this part for compliance demonstrated under 
appendix J of this part. Subsequent civil versions of any such 
helicopter must meet the Stage 2 requirements.
    (d) Helicopters in the primary category:
    (1) Except as provided in paragraph (d)(2) of this section, for a 
helicopter for which application for a type certificate in the primary 
category is made, and that was not previously certificated under 
appendix H of this part, compliance with appendix H of this part must be 
shown.
    (2) For a helicopter that:
    (i) Has a normal or transport type certificate issued under this 
chapter,
    (ii) Has a standard airworthiness certificate issued under this 
chapter,
    (iii) Has not undergone an acoustical change from its type design,
    (iv) Has not previously been certificated under appendix H of this 
part, and
    (v) For which application for conversion to the primary category is 
made, no further showing of compliance with this part is required.

[Doc. No. 26910, 57 FR 42855, Sept. 16, 1992, as amended by Amdt. 36-30, 
79 FR 12045, Mar. 4, 2014]

Subparts I-J [Reserved]



                          Subpart K_Tiltrotors

    Source: 78 FR 1139, Jan. 8, 2013, unless otherwise noted.



Sec. 36.1101  Noise measurement and evaluation.

    For tiltrotors, the noise generated must be measured and evaluated 
under Appendix K of this part, or under an approved equivalent 
procedure.



Sec. 36.1103  Noise limits.

    (a) Compliance with the maximum noise levels prescribed in Appendix 
K of this part must be shown for a tiltrotor for which the application 
for the issuance of a type certificate is made on or after March 11, 
2013.
    (b) To demonstrate compliance with this part, noise levels may not 
exceed the noise limits listed in Appendix K, Section K4, Noise Limits 
of this part. Appendix K of this part (or an approved equivalent 
procedure) must also be used to evaluate and demonstrate compliance with 
the approved test procedures, and at the applicable noise measurement 
points.

Subparts L-N [Reserved]



     Subpart O_Documentation, Operating Limitations and Information



Sec. 36.1501  Procedures, noise levels and other information.

    (a) All procedures, weights, configurations, and other information 
or data employed for obtaining the certified noise levels prescribed by 
this part, including equivalent procedures used for flight, testing, and 
analysis, must be developed and approved. Noise levels achieved during 
type certification must be included in the approved airplane 
(rotorcraft) flight manual.
    (b) Where supplemental test data are approved for modification or 
extension

[[Page 947]]

of an existing flight data base, such as acoustic data from engine 
static tests used in the certification of acoustical changes, the test 
procedures, physical configuration, and other information and procedures 
that are employed for obtaining the supplemental data must be developed 
and approved.

[Amdt. 36-15, 53 FR 16366, May 6, 1988]



Sec. 36.1581  Manuals, markings, and placards.

    (a) If an Airplane Flight Manual or Rotorcraft Flight Manual is 
approved, the approved portion of the Airplane Flight Manual or 
Rotorcraft Flight Manual must contain the following information, in 
addition to that specified under Sec. 36.1583 of this part. If an 
Airplane Flight Manual or Rotorcraft Flight Manual is not approved, the 
procedures and information must be furnished in any combination of 
approved manual material, markings, and placards.
    (1) For transport category large airplanes and jet airplanes, the 
noise level information must be one value for each flyover, lateral, and 
approach as defined and required by appendix B of this part, along with 
the maximum takeoff weight, maximum landing weight, and configuration.
    (2) For propeller driven small airplanes, the noise level 
information must be one value for takeoff as defined and required by 
appendix G of this part, along with the maximum takeoff weight and 
configuration.
    (3) For rotorcraft, the noise level information must be one value 
for each takeoff, flyover, and approach as defined and required by 
appendix H of this part, or one value for flyover as defined and 
required by appendix J of this part, at the maximum takeoff weight and 
configuration.
    (b) If supplemental operational noise level information is included 
in the approved portion of the Airplane Flight Manual, it must be 
segregated, identified as information in addition to the certificated 
noise levels, and clearly distinguished from the information required 
under Sec. 36.1581(a).
    (c) The following statement must be furnished near the listed noise 
levels:

No determination has been made by the Federal Aviation Administration 
that the noise levels of this aircraft are or should be acceptable or 
unacceptable for operation at, into, or out of, any airport.

    (d) For transport category large airplanes and jet airplanes, for 
which the weight used in meeting the takeoff or landing noise 
requirements of this part is less than the maximum weight established 
under the applicable airworthiness requirements, those lesser weights 
must be furnished, as operating limitations in the operating limitations 
section of the Airplane Flight Manual. Further, the maximum takeoff 
weight must not exceed the takeoff weight that is most critical from a 
takeoff noise standpoint.
    (e) For propeller driven small airplanes and for propeller-driven, 
commuter category airplanes for which the weight used in meeting the 
flyover noise requirements of this part is less than the maximum weight 
by an amount exceeding the amount of fuel needed to conduct the test, 
that lesser weight must be furnished, as an operating limitation, in the 
operating limitations section of an approved Airplane Flight Manual, in 
approved manual material, or on an approved placard.
    (f) For primary, normal, transport, and restricted category 
helicopters, if the weight used in meeting the takeoff, flyover, or 
approach noise requirements of appendix H of this part, or the weight 
used in meeting the flyover noise requirement of appendix J of this 
part, is less than the certificated maximum takeoff weight established 
under either Sec. 27.25(a) or Sec. 29.25(a) of this chapter, that 
lesser weight must be furnished as an operating limitation in the 
operating limitations section of the Rotorcraft Flight Manual, in FAA-
approved manual material, or on an FAA-approved placard.
    (g) Except as provided in paragraphs (d), (e), and (f) of this 
section, no operating limitations are furnished under this part.

[Doc. No. 13243, 40 FR 1035, Jan. 6, 1975]

    Editorial Note: For Federal Register citations affecting Sec. 
36.1581, see the List of CFR Sections Affected, which appears in the 
Finding Aids section of the printed volume and at www.fdsys.gov.

[[Page 948]]



Sec. 36.1583  Noncomplying agricultural and fire fighting airplanes.

    (a) This section applies to propeller-driven, small airplanes that--
    (1) Are designed for ``agricultural aircraft operations'' (as 
defined in Sec. 137.3 of this chapter, effective on January 1, 1966) or 
for dispensing fire fighting materials; and
    (2) Have not been shown to comply with the noise levels prescribed 
under appendix F of this part--
    (i) For which application is made for the original issue of a 
standard airworthiness certificate and that do not have any flight time 
before January 1, 1980; or
    (ii) For which application is made for an acoustical change 
approval, for airplanes which have a standard airworthiness certificate 
after the change in the type design, and that do not have any flight 
time in the changed configuration before January 1, 1980.
    (b) For airplanes covered by this section an operating limitation 
reading as follows must be furnished in the manner prescribed in Sec. 
36.1581:

    Noise abatement: This airplane has not been shown to comply with the 
noise limits in FAR Part 36 and must be operated in accordance with the 
noise operating limitation prescribed under FAR Sec. 91.815.

[Amdt. 36-11, 45 FR 67066, Oct. 9, 1980. Redesignated by Amdt. 36-14, 53 
FR 3540, Feb. 5, 1988; Amdt. 36-18, 54 FR 34330, Aug. 18, 1989]



 Sec. Appendix A to Part 36--Aircraft Noise Measurement and Evaluation 
                           Under Sec. 36.101

Sec.

A36.1 Introduction.
A36.2 Noise Certification Test and Measurement Conditions.
A36.3 Measurement of Airplane Noise Received on the Ground.
A36.4 Calculations of Effective Perceived Noise Level From Measured 
          Data.
A36.5 Reporting of Data to the FAA.
A36.6 Nomenclature: Symbols and Units.
A36.7 Sound Attenuation in Air.
A36.8 [Reserved]
A36.9 Adjustment of Airplane Flight Test Results.

                       Section A36.1 Introduction

    A36.1.1 This appendix prescribes the conditions under which airplane 
noise certification tests must be conducted and states the measurement 
procedures that must be used to measure airplane noise. The procedures 
that must be used to determine the noise evaluation quantity designated 
as effective perceived noise level, EPNL, under Sec. Sec. 36.101 and 
36.803 are also stated.
    A36.1.2 The instructions and procedures given are intended to ensure 
uniformity during compliance tests and to permit comparison between 
tests of various types of airplanes conducted in various geographical 
locations.
    A36.1.3 A complete list of symbols and units, the mathematical 
formulation of perceived noisiness, a procedure for determining 
atmospheric attenuation of sound, and detailed procedures for correcting 
noise levels from non-reference to reference conditions are included in 
this appendix.
    A36.1.4 For Stage 4 airplanes, an acceptable alternate for noise 
measurement and evaluation is Appendix 2 to the International Civil 
Aviation Organization (ICAO) Annex 16, Environmental Protection, Volume 
I, Aircraft Noise, Third Edition, July 1993, Amendment 7, effective 
March 21, 2002. [Incorporated by reference, see Sec. 36.6].

    Section A36.2 Noise Certification Test and Measurement Conditions

    A36.2.1 General.
    A36.2.1.1 This section prescribes the conditions under which noise 
certification must be conducted and the measurement procedures that must 
be used.

    Note: Many noise certifications involve only minor changes to the 
airplane type design. The resulting changes in noise can often be 
established reliably without resorting to a complete test as outlined in 
this appendix. For this reason, the FAA permits the use of approved 
equivalent procedures. There are also equivalent procedures that may be 
used in full certification tests, in the interest of reducing costs and 
providing reliable results. Guidance material on the use of equivalent 
procedures in the noise certification of subsonic jet and propeller-
driven large airplanes is provided in the current advisory circular for 
this part.

    A36.2.2 Test environment.
    A36.2.2.1 Locations for measuring noise from an airplane in flight 
must be surrounded by relatively flat terrain having no excessive sound 
absorption characteristics such as might be caused by thick, matted, or 
tall grass, shrubs, or wooded areas. No obstructions that significantly 
influence the sound field from the airplane must exist within a conical 
space above the point on the ground vertically below the microphone, the 
cone being defined by an axis normal to the ground and by a half-angle 
80[deg] from this axis.

    Note: Those people carrying out the measurements could themselves 
constitute such obstruction.


[[Page 949]]


    A36.2.2.2 The tests must be carried out under the following 
atmospheric conditions.
    (a) No precipitation;
    (b) Ambient air temperature not above 95 [deg]F (35 [deg]C) and not 
below 14 [deg]F (-10 [deg]C), and relative humidity not above 95% and 
not below 20% over the whole noise path between a point 33 ft (10 m) 
above the ground and the airplane;

    Note: Care should be taken to ensure that the noise measuring, 
airplane flight path tracking, and meteorological instrumentation are 
also operated within their specific environmental limitations.

    (c) Relative humidity and ambient temperature over the whole noise 
path between a point 33 ft (10 m) above the ground and the airplane such 
that the sound attenuation in the one-third octave band centered on 8 
kHz will not be more than 12 dB/100 m unless:
    (1) The dew point and dry bulb temperatures are measured with a 
device which is accurate to 0.9 [deg]F (0.5 [deg]C) and used to obtain relative humidity; in 
addition layered sections of the atmosphere are used as described in 
section A36.2.2.3 to compute equivalent weighted sound attenuations in 
each one-third octave band; or
    (2) The peak noy values at the time of PNLT, after adjustment to 
reference conditions, occur at frequencies less than or equal to 400 
Hz.;
    (d) If the atmospheric absorption coefficients vary over the PNLTM 
sound propagation path by more than 1.6 dB/1000 ft 
(0.5 dB/100m) in the 3150Hz one-third octave band 
from the value of the absorption coefficient derived from the 
meteorological measurement obtained at 33 ft (10 m) above the surface, 
``layered'' sections of the atmosphere must be used as described in 
section A36.2.2.3 to compute equivalent weighted sound attenuations in 
each one-third octave band; the FAA will determine whether a sufficient 
number of layered sections have been used. For each measurement, where 
multiple layering is not required, equivalent sound attenuations in each 
one-third octave band must be determined by averaging the atmospheric 
absorption coefficients for each such band at 33 ft (10 m) above ground 
level, and at the flight level of the airplane at the time of PNLTM, for 
each measurement;
    (e) Average wind velocity 33 ft (10 m) above ground may not exceed 
12 knots and the crosswind velocity for the airplane may not exceed 7 
knots. The average wind velocity must be determined using a 30-second 
averaging period spanning the 10 dB-down time interval. Maximum wind 
velocity 33 ft (10 m) above ground is not to exceed 15 knots and the 
crosswind velocity is not to exceed 10 knots during the 10 dB-down time 
interval;
    (f) No anomalous meteorological or wind conditions that would 
significantly affect the measured noise levels when the noise is 
recorded at the measuring points specified by the FAA; and
    (g) Meteorological measurements must be obtained within 30 minutes 
of each noise test measurement; meteorological data must be interpolated 
to actual times of each noise measurement.
    A36.2.2.3 When a multiple layering calculation is required by 
section A36.2.2.2(c) or A36.2.2.2(d) the atmosphere between the airplane 
and 33 ft (10 m) above the ground must be divided into layers of equal 
depth. The depth of the layers must be set to not more than the depth of 
the narrowest layer across which the variation in the atmospheric 
absorption coefficient of the 3150 Hz one-third octave band is not 
greater than 1.6 dB/1000 ft (0.5 dB/100m), with a minimum layer depth of 100 ft (30 
m). This requirement must be met for the propagation path at PNLTM. The 
mean of the values of the atmospheric absorption coefficients at the top 
and bottom of each layer may be used to characterize the absorption 
properties of each layer.
    A36.2.2.4 The airport control tower or another facility must be 
aproved by the FAA for use as the central location at which measurements 
of atmospheric parameters are representative of those conditions 
existing over the geographical area in which noise measurements are 
made.
    A36.2.3 Flight path measurement.
    A36.2.3.1 The airplane height and lateral position relative to the 
flight track must be determined by a method independent of normal flight 
instrumentation such as radar tracking, theodolite triangulation, or 
photographic scaling techniques, to be approved by the FAA.
    A36.2.3.2 The airplane position along the flight path must be 
related to the noise recorded at the noise measurement locations by 
means of synchronizing signals over a distance sufficient to assure 
adequate data during the period that the noise is within 10 dB of the 
maximum value of PNLT.
    A36.2.3.3 Position and performance data required to make the 
adjustments referred to in section A36.9 of this appendix must be 
automatically recorded at an approved sampling rate. Measuring equipment 
must be approved by the FAA.

   Section A36.3 Measurement of Airplane Noise Received on the Ground

    A36.3.1 Definitions.
    For the purposes of section A36.3 the following definitions apply:
    A36.3.1.1 Measurement system means the combination of instruments 
used for the measurement of sound pressure levels, including a sound 
calibrator, windscreen, microphone system, signal recording and 
conditioning devices, and one-third octave band analysis system.


[[Page 950]]


    Note: Practical installations may include a number of microphone 
systems, the outputs from which are recorded simultaneously by a multi-
channel recording/analysis device via signal conditioners, as 
appropriate. For the purpose of this section, each complete measurement 
channel is considered to be a measurement system to which the 
requirements apply accordingly.

    A36.3.1.2 Microphone system means the components of the measurement 
system which produce an electrical output signal in response to a sound 
pressure input signal, and which generally include a microphone, a 
preamplifier, extension cables, and other devices as necessary.
    A36.3.1.3 Sound incidence angle means in degrees, an angle between 
the principal axis of the microphone, as defined in IEC 61094-3 and IEC 
61094-4, as amended and a line from the sound source to the center of 
the diaphragm of the microphone.

    Note: When the sound incidence angle is 0[deg], the sound is said to 
be received at the microphone at ``normal (perpendicular) incidence;'' 
when the sound incidence angle is 90[deg], the sound is said to be 
received at ``grazing incidence.''

    A36.3.1.4 Reference direction means, in degrees, the direction of 
sound incidence specified by the manufacturer of the microphone, 
relative to a sound incidence angle of 0[deg], for which the free-field 
sensitivity level of the microphone system is within specified tolerance 
limits.
    A36.3.1.5 Free-field sensitivity of a microphone system means, in 
volts per Pascal, for a sinusoidal plane progressive sound wave of 
specified frequency, at a specified sound incidence angle, the quotient 
of the root mean square voltage at the output of a microphone system and 
the root mean square sound pressure that would exist at the position of 
the microphone in its absence.
    A36.3.1.6 Free-field sensitivity level of a microphone system means, 
in decibels, twenty times the logarithm to the base ten of the ratio of 
the free-field sensitivity of a microphone system and the reference 
sensitivity of one volt per Pascal.

    Note: The free-field sensitivity level of a microphone system may be 
determined by subtracting the sound pressure level (in decibels re 20 
[micro]Pa) of the sound incident on the microphone from the voltage 
level (in decibels re 1 V) at the output of the microphone system, and 
adding 93.98 dB to the result.

    A36.3.1.7 Time-average band sound pressure level means in decibels, 
ten times the logarithm to the base ten, of the ratio of the time mean 
square of the instantaneous sound pressure during a stated time interval 
and in a specified one-third octave band, to the square of the reference 
sound pressure of 20 [micro]Pa.
    A36.3.1.8 Level range means, in decibels, an operating range 
determined by the setting of the controls that are provided in a 
measurement system for the recording and one-third octave band analysis 
of a sound pressure signal. The upper boundary associated with any 
particular level range must be rounded to the nearest decibel.
    A36.3.1.9 Calibration sound pressure level means, in decibels, the 
sound pressure level produced, under reference environmental conditions, 
in the cavity of the coupler of the sound calibrator that is used to 
determine the overall acoustical sensitivity of a measurement system.
    A36.3.1.10 Reference level range means, in decibels, the level range 
for determining the acoustical sensitivity of the measurement system and 
containing the calibration sound pressure level.
    A36.3.1.11 Calibration check frequency means, in hertz, the nominal 
frequency of the sinusoidal sound pressure signal produced by the sound 
calibrator.
    A36.3.1.12 Level difference means, in decibels, for any nominal one-
third octave midband frequency, the output signal level measured on any 
level range minus the level of the corresponding electrical input 
signal.
    A36.3.1.13 Reference level difference means, in decibels, for a 
stated frequency, the level difference measured on a level range for an 
electrical input signal corresponding to the calibration sound pressure 
level, adjusted as appropriate, for the level range.
    A36.3.1.14 Level non-linearity means, in decibels, the level 
difference measured on any level range, at a stated one-third octave 
nominal midband frequency, minus the corresponding reference level 
difference, all input and output signals being relative to the same 
reference quantity.
    A36.3.1.15 Linear operating range means, in decibels, for a stated 
level range and frequency, the range of levels of steady sinusoidal 
electrical signals applied to the input of the entire measurement 
system, exclusive of the microphone but including the microphone 
preamplifier and any other signal-conditioning elements that are 
considered to be part of the microphone system, extending from a lower 
to an upper boundary, over which the level non-linearity is within 
specified tolerance limits.

    Note: Microphone extension cables as configured in the field need 
not be included for the linear operating range determination.

    A36.3.1.16 Windscreen insertion loss means, in decibels, at a stated 
nominal one-third octave midband frequency, and for a stated sound 
incidence angle on the inserted microphone, the indicated sound pressure 
level without the windscreen installed around the microphone minus the 
sound pressure level with the windscreen installed.
    A36.3.2 Reference environmental conditions.

[[Page 951]]

    A36.3.2.1 The reference environmental conditions for specifying the 
performance of a measurement system are:
    (a) Air temperature 73.4 [deg]F (23 [deg]C);
    (b) Static air pressure 101.325 kPa; and
    (c) Relative humidity 50%.
    A36.3.3. General.

    Note: Measurements of aircraft noise that are made using instruments 
that conform to the specifications of this section will yield one-third 
octave band sound pressure levels as a function of time. These one-third 
octave band levels are to be used for the calculation of effective 
perceived noise level as described in section A36.4.

    A36.3.3.1 The measurement system must consist of equipment approved 
by the FAA and equivalent to the following:
    (a) A windscreen (See A36.3.4.);
    (b) A microphone system (See A36.3.5):
    (c) A recording and reproducing system to store the measured 
aircraft noise signals for subsequent analysis (see A36.3.6);
    (d) A one-third octave band analysis system (see A36.3.7); and
    (e) Calibration systems to maintain the acoustical sensitivity of 
the above systems within specified tolerance limits (see A36.3.8).
    A36.3.3.2. For any component of the measurement system that converts 
an analog signal to digital form, such conversion must be performed so 
that the levels of any possible aliases or artifacts of the digitization 
process will be less than the upper boundary of the linear operating 
range by at least 50 dB at any frequency less than 12.5 kHz. The 
sampling rate must be at least 28 kHz. An anti-aliasing filter must be 
included before the digitization process.
    A36.3.4 Windscreen.
    A36.3.4.1 In the absence of wind and for sinusoidal sounds at 
grazing incidence, the insertion loss caused by the windscreen of a 
stated type installed around the microphone must not exceed 1.5 dB at nominal one-third octave midband frequencies 
from 50 Hz to 10 kHz inclusive.
    A36.3.5 Microphone system.
    A36.3.5.1 The microphone system must meet the specifications in 
sections A36.3.5.2 to A36.3.5.4. Various microphone systems may be 
approved by the FAA on the basis of demonstrated equivalent overall 
electroacoustical performance. Where two or more microphone systems of 
the same type are used, demonstration that at least one system conforms 
to the specifications in full is sufficient to demonstrate conformance.

    Note: An applicant must still calibrate and check each system as 
required in section A36.3.9.

    A36.3.5.2 The microphone must be mounted with the sensing element 4 
ft (1.2 m) above the local ground surface and must be oriented for 
grazing incidence, i.e., with the sensing element substantially in the 
plane defined by the predicted reference flight path of the aircraft and 
the measuring station. The microphone mounting arrangement must minimize 
the interference of the supports with the sound to be measured. Figure 
A36-1 illustrates sound incidence angles on a microphone.
    A36.3.5.3 The free-field sensitivity level of the microphone and 
preamplifier in the reference direction, at frequencies over at least 
the range of one-third-octave nominal midband frequencies from 50 Hz to 
5 kHz inclusive, must be within 1.0 dB of that at 
the calibration check frequency, and within 2.0 dB 
for nominal midband frequencies of 6.3 kHz, 8 kHz and 10 kHz.
    A36.3.5.4 For sinusoidal sound waves at each one-third octave 
nominal midband frequency over the range from 50 Hz to 10 kHz inclusive, 
the free-field sensitivity levels of the microphone system at sound 
incidence angles of 30[deg], 60[deg], 90[deg], 120[deg] and 150[deg], 
must not differ from the free-field sensitivity level at a sound 
incidence angle of 0[deg] (``normal incidence'') by more than the values 
shown in Table A36-1. The free-field sensitivity level differences at 
sound incidence angles between any two adjacent sound incidence angles 
in Table A36-1 must not exceed the tolerance limit for the greater 
angle.

[[Page 952]]

[GRAPHIC] [TIFF OMITTED] TR08JY02.000

    A36.3.6 Recording and reproducing systems.
    A36.3.6.1 A recording and reproducing system, such as a digital or 
analog magnetic tape recorder, a computer-based system or other 
permanent data storage device, must be used to store sound pressure 
signals for subsequent analysis. The sound produced by the aircraft must 
be recorded in such a way that a record of the complete acoustical 
signal is retained. The recording and reproducing systems must meet the 
specifications in sections A36.3.6.2 to A36.3.6.9 at the recording 
speeds and/or data sampling rates used for the noise certification 
tests. Conformance must be demonstrated for the frequency bandwidths and 
recording channels selected for the tests.
    A36.3.6.2 The recording and reproducing systems must be calibrated 
as described in section A36.3.9.
    (a) For aircraft noise signals for which the high frequency spectral 
levels decrease rapidly with increasing frequency, appropriate pre-
emphasis and complementary de-emphasis networks may be included in the 
measurement system. If pre-emphasis is included, over the range of 
nominal one-third octave midband frequencies from 800 Hz to 10 kHz 
inclusive, the electrical gain provided by the pre-emphasis network must 
not exceed 20 dB relative to the gain at 800 Hz.
    A36.3.6.3 For steady sinusoidal electrical signals applied to the 
input of the entire measurement system including all parts of the 
microphone system except the microphone at a selected signal level 
within 5 dB of that corresponding to the calibration sound pressure 
level on the reference level range, the time-average signal level 
indicated by the readout device at any one-third octave nominal midband 
frequency from 50 Hz to 10 kHz inclusive must be within 1.5 dB of that at the calibration check frequency. The 
frequency response of a measurement system, which includes components 
that convert analog signals to digital form, must be within 0.3 dB of the response at 10 kHz over the frequency 
range from 10 kHz to 11.2 kHz.

    Note: Microphone extension cables as configured in the field need 
not be included for the frequency response determination. This

[[Page 953]]

allowance does not eliminate the requirement of including microphone 
extension cables when performing the pink noise recording in section 
A36.3.9.5.

    A36.3.6.4 For analog tape recordings, the amplitude fluctuations of 
a 1 kHz sinusoidal signal recorded within 5 dB of the level 
corresponding to the calibration sound pressure level must not vary by 
more than 0.5 dB throughout any reel of the type 
of magnetic tape used. Conformance to this requirement must be 
demonstrated using a device that has time-averaging properties 
equivalent to those of the spectrum analyzer.
    A36.3.6.5 For all appropriate level ranges and for steady sinusoidal 
electrical signals applied to the input of the measurement system, 
including all parts of the microphone system except the microphone, at 
one-third-octave nominal midband frequencies of 50 Hz, 1 kHz and 10 kHz, 
and the calibration check frequency, if it is not one of these 
frequencies, the level non-linearity must not exceed 0.5 dB for a linear operating range of at least 50 dB 
below the upper boundary of the level range.

    Note 1: Level linearity of measurement system components may be 
tested according to the methods described in IEC 61265 as amended.
    Note 2: Microphone extension cables configured in the field need not 
be included for the level linearity determination.

    A36.3.6.6 On the reference level range, the level corresonding to 
the calibration sound pressure level must be at least 5 dB, but no more 
than 30 dB less than the upper boundary of the level range.
    A36.3.6.7 The linear operating ranges on adjacent level ranges must 
overlap by at least 50 dB minus the change in attenuation introduced by 
a change in the level range controls.

    Note: It is possible for a measurement system to have level range 
controls that permit attenuation changes of either 10 dB or 1 dB, for 
example. With 10 dB steps, the minimum overlap required would be 40 dB, 
and with 1 dB steps the minimum overlap would be 49 dB.

    A36.3.6.8 An overload indicator must be included in the recording 
and reproducing systems so that an overload indication will occur during 
an overload condition on any relevant level range.
    A36.3.6.9 Attenuators included in the measurement system to permit 
range changes must operate in known intervals of decibel steps.
    A36.3.7 Analysis systems.
    A36.3.7.1 The analysis system must conform to the specifications in 
sections A36.3.7.2 to A36.3.7.7 for the frequency bandwidths, channel 
configurations and gain settings used for analysis.
    A36.3.7.2 The output of the analysis system must consist of one-
third octave band sound pressure levels as a function of time, obtained 
by processing the noise signals (preferably recorded) through an 
analysis system with the following characteristics:
    (a) A set of 24 one-third octave band filters, or their equivalent, 
having nominal midband frequencies from 50 Hz to 10 kHz inclusive;
    (b) Response and averaging properties in which, in principle, the 
output from any one-third octave filter band is squared, averaged and 
displayed or stored as time-averaged sound pressure levels;
    (c) The interval between successive sound pressure level samples 
must be 500 ms 5 milliseconds(ms) for spectral 
analysis with or without slow time-weighting, as defined in section 
A36.3.7.4;
    (d) For those analysis systems that do not process the sound 
pressure signals during the period of time required for readout and/or 
resetting of the analyzer, the loss of data must not exceed a duration 
of 5 ms; and
    (e) The analysis system must operate in real time from 50 Hz through 
at least 12 kHz inclusive. This requirement applies to all operating 
channels of a multi-channel spectral analysis system.
    A36.3.7.3 The minimum standard for the one-third octave band 
analysis system is the class 2 electrical performance requirements of 
IEC 61260 as amended, over the range of one-third octave nominal midband 
frequencies from 50 Hz through 10 kHz inclusive.

    Note: IEC 61260 specifies procedures for testing of one-third octave 
band analysis systems for relative attenuation, anti-aliasing filters, 
real time operation, level linearity, and filter integrated response 
(effective bandwidth).

    A36.3.7.4 When slow time averaging is performed in the analyzer, the 
response of the one-third octave band analysis system to a sudden onset 
or interruption of a constant sinusoidal signal at the respective one-
third octave nominal midband frequency, must be measured at sampling 
instants 0.5, 1, 1.5 and 2 seconds(s) after the onset and 0.5 and 1s 
after interruption. The rising response must be -4 1 dB at 0.5s, -1.75 0.75 dB at 1s, 
-1 0.5 dB at 1.5s and -0.5 0.5 dB at 2s relative to the steady-state level. The 
falling response must be such that the sum of the output signal levels, 
relative to the initial steady-state level, and the corresponding rising 
response reading is -6.5 1 dB, at both 0.5 and 1s. 
At subsequent times the sum of the rising and falling responses must be 
-7.5 dB or less. This equates to an exponential averaging process (slow 
time-weighting) with a nominal 1s time constant (i.e., 2s averaging 
time).
    A36.3.7.5 When the one-third octave band sound pressure levels are 
determined from

[[Page 954]]

the output of the analyzer without slow time-weighting, slow time-
weighting must be simulated in the subsequent processing. Simulated slow 
time-weighted sound pressure levels can be obtained using a continuous 
exponential averaging process by the following equation:

Ls (i,k)=10 log [(0.60653) 100.1 Ls[i, (k-1)] + 
          (0.39347) 100.1 L (i, k)]

where Ls(i,k) is the simulated slow time-weighted sound 
          pressure level and L(i,k) is the as-measured 0.5s time average 
          sound pressure level determined from the output of the 
          analyzer for the k-th instant of time and i-th one-third 
          octave band. For k=1, the slow time-weighted sound pressure 
          Ls[i, (k-1=0)] on the right hand side should be set 
          to 0 dB. An approximation of the continuous exponential 
          averaging is represented by the following equation for a four 
          sample averaging process for k =4:

Ls (i,k)=10 log [(0.13) 100.1 L[i,(k-3)] + (0.21) 
          100.1 L[i, (k-2)] + (0.27) 
          100.1 L[i, (k-1)] + (0.39) 
          100.1 L[i, k]]

where Ls (i, k) is the simulated slow time-weighted sound 
          pressure level and L (i, k) is the as measured 0.5s time 
          average sound pressure level determined from the output of the 
          analyzer for the k-th instant of time and the i-th one-third 
          octave band.
    The sum of the weighting factors is 1.0 in the two equations. Sound 
pressure levels calculated by means of either equation are valid for the 
sixth and subsequent 0.5s data samples, or for times greater than 2.5s 
after initiation of data analysis.

    Note: The coefficients in the two equations were calculated for use 
in determining equivalent slow time-weighted sound pressure levels from 
samples of 0.5s time average sound pressure levels. The equations do not 
work with data samples where the averaging time differs from 0.5s.

    A36.3.7.6 The instant in time by which a slow time-weighted sound 
pressure level is characterized must be 0.75s earlier than the actual 
readout time.

    Note: The definition of this instant in time is needed to correlate 
the recorded noise with the aircraft position when the noise was emitted 
and takes into account the averaging period of the slow time-weighting. 
For each 0.5 second data record this instant in time may also be 
identified as 1.25 seconds after the start of the associated 2 second 
averaging period.

    A36.3.7.7 The resolution of the sound pressure levels, both 
displayed and stored, must be 0.1 dB or finer.
    A36.3.8 Calibration systems.
    A36.3.8.1 The acoustical sensitivity of the measurement system must 
be determined using a sound calibrator generating a known sound pressure 
level at a known frequency. The minimum standard for the sound 
calibrator is the class 1L requirements of IEC 60942 as amended.
    A36.3.9 Calibration and checking of system.
    A36.3.9.1 Calibration and checking of the measurement system and its 
constituent components must be carried out to the satisfaction of the 
FAA by the methods specified in sections A36.3.9.2 through A36.3.9.10. 
The calibration adjustments, including those for environmental effects 
on sound calibrator output level, must be reported to the FAA and 
applied to the measured one-third-octave sound pressure levels 
determined from the output of the analyzer. Data collected during an 
overload indication are invalid and may not be used. If the overload 
condition occurred during recording, the associated test data are 
invalid, whereas if the overload occurred during analysis, the analysis 
must be repeated with reduced sensitivity to eliminate the overload.
    A36.3.9.2 The free-field frequency response of the microphone system 
may be determined by use of an electrostatic actuator in combination 
with manufacturer's data or by tests in an anechoic free-field facility. 
The correction for frequency response must be determined within 90 days 
of each test series. The correction for non-uniform frequency response 
of the microphone system must be reported to the FAA and applied to the 
measured one-third octave band sound pressure levels determined from the 
output of the analyzer.
    A36.3.9.3 When the angles of incidence of sound emitted from the 
aircraft are within 30[deg] of grazing incidence 
at the microphone (see Figure A36-1), a single set of free-field 
corrections based on grazing incidence is considered sufficient for 
correction of directional response effects. For other cases, the angle 
of incidence for each 0.5 second sample must be determined and applied 
for the correction of incidence effects.
    A36.3.9.4 For analog magnetic tape recorders, each reel of magnetic 
tape must carry at least 30 seconds of pink random or pseudo-random 
noise at its beginning and end. Data obtained from analog tape-recorded 
signals will be accepted as reliable only if level differences in the 10 
kHz one-third-octave-band are not more than 0.75 dB for the signals 
recorded at the beginning and end.
    A36.3.9.5 The frequency response of the entire measurement system 
while deployed in the field during the test series, exclusive of the 
microphone, must be determined at a level within 5 dB of the level 
corresponding to the calibration sound pressure level on the level range 
used during the tests for each one-third octave nominal midband 
frequency from 50 Hz to 10 kHz inclusive, utilizing pink random or 
pseudo-random noise. Within six months of each test series the output of 
the noise generator must be determined by a

[[Page 955]]

method traceable to the U.S. National Institute of Standards and 
Technology or to an equivalent national standards laboratory as 
determined by the FAA. Changes in the relative output from the previous 
calibration at each one-third octave band may not exceed 0.2 dB. The 
correction for frequency response must be reported to the FAA and 
applied to the measured one-third octave sound pressure levels 
determined from the output of the analyzer.
    A36.3.9.6 The performance of switched attenuators in the equipment 
used during noise certification measurements and calibration must be 
checked within six months of each test series to ensure that the maximum 
error does not exceed 0.1 dB.
    A36.3.9.7 The sound pressure level produced in the cavity of the 
coupler of the sound calibrator must be calculated for the test 
environmental conditions using the manufacturer's supplied information 
on the influence of atmospheric air pressure and temperature. This sound 
pressure level is used to establish the acoustical sensitivity of the 
measurement system. Within six months of each test series the output of 
the sound calibrator must be determined by a method traceable to the 
U.S. National Institute of Standards and Technology or to an equivalent 
national standards laboratory as determined by the FAA. Changes in 
output from the previous calibration must not exceed 0.2 dB.
    A36.3.9.8 Sufficient sound pressure level calibrations must be made 
during each test day to ensure that the acoustical sensitivity of the 
measurement system is known at the prevailing environmental conditions 
corresponding with each test series. The difference between the 
acoustical sensitivity levels recorded immediately before and 
immediately after each test series on each day may not exceed 0.5 dB. 
The 0.5 dB limit applies after any atmospheric pressure corrections have 
been determined for the calibrator output level. The arithmetic mean of 
the before and after measurements must be used to represent the 
acoustical sensitivity level of the measurement system for that test 
series. The calibration corrections must be reported to the FAA and 
applied to the measured one-third octave band sound pressure levels 
determined from the output of the analyzer.
    A36.3.9.9 Each recording medium, such as a reel, cartridge, 
cassette, or diskette, must carry a sound pressure level calibration of 
at least 10 seconds duration at its beginning and end.
    A36.3.9.10 The free-field insertion loss of the windscreen for each 
one-third octave nominal midband frequency from 50 Hz to 10 kHz 
inclusive must be determined with sinusoidal sound signals at the 
incidence angles determined to be applicable for correction of 
directional response effects per section A36.3.9.3. The interval between 
angles tested must not exceed 30 degrees. For a windscreen that is 
undamaged and uncontaminated, the insertion loss may be taken from 
manufacturer's data. Alternatively, within six months of each test 
series the insertion loss of the windscreen may be determined by a 
method traceable to the U.S. National Institute of Standards and 
Technology or an equivalent national standards laboratory as determined 
by the FAA. Changes in the insertion loss from the previous calibration 
at each one-third-octave frequency band must not exceed 0.4 dB. The 
correction for the free-field insertion loss of the windscreen must be 
reported to the FAA and applied to the measured one-third octave sound 
pressure levels determined from the output of the analyzer.
    A36.3.10 Adjustments for ambient noise.
    A36.3.10.1 Ambient noise, including both an acoustical background 
and electrical noise of the measurement system, must be recorded for at 
least 10 seconds at the measurement points with the system gain set at 
the levels used for the aircraft noise measurements. Ambient noise must 
be representative of the acoustical background that exists during the 
flyover test run. The recorded aircraft noise data is acceptable only if 
the ambient noise levels, when analyzed in the same way, and quoted in 
PNL (see A36.4.1.3 (a)), are at least 20 dB below the maximum PNL of the 
aircraft.
    A36.3.10.2 Aircraft sound pressure levels within the 10 dB-down 
points (see A36.4.5.1) must exceed the mean ambient noise levels 
determined in section A36.3.10.1 by at least 3 dB in each one-third 
octave band, or must be adjusted using a method approved by the FAA; one 
method is described in the current advisory circular for this part.

   Section A36.4 Calculation of Effective Perceived Noise Level From 
                              Measured Data

    A36.4.1 General.
    A36.4.1.1 The basic element for noise certification criteria is the 
noise evaluation measure known as effective perceived noise level, EPNL, 
in units of EPNdB, which is a single number evaluator of the subjective 
effects of airplane noise on human beings. EPNL consists of 
instantaneous perceived noise level, PNL, corrected for spectral 
irregularities, and for duration. The spectral irregularity correction, 
called ``tone correction factor'', is made at each time increment for 
only the maximum tone.
    A36.4.1.2 Three basic physical properties of sound pressure must be 
measured: level, frequency distribution, and time variation. To 
determine EPNL, the instantaneous sound pressure level in each of the 24 
one-third octave bands is required for each 0.5 second increment of time 
during the airplane noise measurement.

[[Page 956]]

    A36.4.1.3 The calculation procedure that uses physical measurements 
of noise to derive the EPNL evaluation measure of subjective response 
consists of the following five steps:
    (a) The 24 one-third octave bands of sound pressure level are 
converted to perceived noisiness (noy) using the method described in 
section A36.4.2.1 (a). The noy values are combined and then converted to 
instantaneous perceived noise levels, PNL(k).
    (b) A tone correction factor C(k) is calculated for each spectrum to 
account for the subjective response to the presence of spectral 
irregularities.
    (c) The tone correction factor is added to the perceived noise level 
to obtain tone-corrected perceived noise levels PNLT(k), at each one-
half second increment:

PNLT(k)=PNL(k) + C(k)

The instantaneous values of tone-corrected perceived noise level are 
derived and the maximum value, PNLTM, is determined.
    (d) A duration correction factor, D, is computed by integration 
under the curve of tone-corrected perceived noise level versus time.
    (e) Effective perceived noise level, EPNL, is determined by the 
algebraic sum of the maximum tone-corrected perceived noise level and 
the duration correction factor:

EPNL=PNLTM + D

    A36.4.2 Perceived noise level.
    A36.4.2.1 Instantaneous perceived noise levels, PNL(k), must be 
calculated from instantaneous one-third octave band sound pressure 
levels, SPL(i, k) as follows:
    (a) Step 1: For each one-third octave band from 50 through 10,000 
Hz, convert SPL(i, k) to perceived noisiness n(i, k), by using the 
mathematical formulation of the noy table given in section A36.4.7.
    (b) Step 2: Combine the perceived noisiness values, n(i, k), 
determined in step 1 by using the following formula:
[GRAPHIC] [TIFF OMITTED] TR08JY02.001

where n(k) is the largest of the 24 values of n(i, k) and N(k) is the 
          total perceived noisiness.
    (c) Step 3: Convert the total perceived noisiness, N(k), determined 
in Step 2 into perceived noise level, PNL(k), using the following 
formula:
[GRAPHIC] [TIFF OMITTED] TR08JY02.002

    Note: PNL(k) is plotted in the current advisory circular for this 
part.
    A36.4.3 Correction for spectral irregularities.

    A36.4.3.1 Noise having pronounced spectral irregularities (for 
example, the maximum discrete frequency components or tones) must be 
adjusted by the correction factor C(k) calculated as follows:
    (a) Step 1: After applying the corrections specified under section 
A36.3.9, start with the sound pressure level in the 80 Hz one-third 
octave band (band number 3), calculate the changes in sound pressure 
level (or ``slopes'') in the remainder of the one-third octave bands as 
follows:

s(3,k)=no value
s(4,k)=SPL(4,k)-SPL(3,k)


s(i,k)=SPL(i,k)-SPL(i-1,k)


s(24,k)=SPL(24,k)-SPL(23,k)

    (b) Step 2: Encircle the value of the slope, s(i, k), where the 
absolute value of the change in slope is greater than five; that is 
where:

[verbar][Delta]s(i,k)[verbar]=[verbar]s(i,k)-s(i-
          1,k)[verbar]5

    (c) Step 3:
    (1) If the encircled value of the slope s(i, k) is positive and 
algebraically greater than the slope s(i-1, k) encircle SPL(i, k).
    (2) If the encircled value of the slope s(i, k) is zero or negative 
and the slope s(i-1, k) is positive, encircle SPL(i-1, k).
    (3) For all other cases, no sound pressure level value is to be 
encircled.
    (d) Step 4: Compute new adjusted sound pressure levels SPL'(i, k) as 
follows:
    (1) For non-encircled sound pressure levels, set the new sound 
pressure levels equal to the original sound pressure levels, SPL'(i, 
k)=SPL(i, k).
    (2) For encircled sound pressure levels in bands 1 through 23 
inclusive, set the new sound pressure level equal to the arithmetic 
average of the preceding and following sound pressure levels as shown 
below:

SPL'(i,k)=\1/2\[SPL(i-1,k)+SPL(i+1,k)]

    (3) If the sound pressure level in the highest frequency band (i=24) 
is encircled, set the new sound pressure level in that band equal to:

SPL'(24,k)=SPL(23,k)+s(23,k)

    (e) Step 5: Recompute new slope s'(i, k), including one for an 
imaginary 25th band, as follows:

s'(3,k)=s'(4,k)
s'(4,k)=SPL'(4,k)-SPL'(3,k)


s'(i,k)=SPL'(i,k)-SPL'(i-1,k)



[[Page 957]]

s'(24,k)=SPL'(24,k)-SPL'(23,k)
s'(25,k)=s'(24,k)

    (f) Step 6: For i, from 3 through 23, compute the arithmetic average 
of the three adjacent slopes as follows:

s(i,k)=\1/3\[s'(i,k)+s'(i+1,k)+s'(i+2,k)]

    (g) Step 7: Compute final one-third octave-band sound pressure 
levels, SPL' (i,k), by beginning with band number 3 and proceeding to 
band number 24 as follows:

    SPL'(3,k)=SPL(3,k)
    SPL'(4,k)=SPL'(3,k)+s(3,k)
    
    
    SPL'(i,k)=SPL'(i-1,k)+s(i-1,k)
    
    
    SPL'(24,k)=SPL'(23,k)+s(23,k)

    (h) Setp 8: Calculate the differences, F (i,k), between the original 
sound pressure level and the final background sound pressure level as 
follows:
F(i,k)=SPL(i,k)-SPL'(i,k)

and note only values equal to or greater than 1.5.
    (i) Step 9: For each of the relevant one-third octave bands (3 
through 24), determine tone correction factors from the sound pressure 
level differences F (i, k) and Table A36-2.

[[Page 958]]

[GRAPHIC] [TIFF OMITTED] TR08JY02.003

    (j) Step 10: Designate the largest of the tone correction factors, 
determined in Step 9, as C(k). (An example of the tone correction 
procedure is given in the current advisory circular for this part). 
Tone-corrected perceived noise levels PNLT(k) must be determined by 
adding the C(k) values to corresponding PNL(k) values, that is:

PNLT(k)=PNL(k)+C(k)

For any i-th one-third octave band, at any k-th increment of time, for 
which the tone correction factor is suspected to result from something 
other than (or in addition to) an actual tone (or any spectral 
irregularity other than airplane noise), an additional analysis may be 
made using a filter with a bandwidth narrower than one-third of an 
octave. If the narrow band analysis corroborates these suspicions, then 
a revised value for the background sound pressure level

[[Page 959]]

SPL'(i,k), may be determined from the narrow band analysis and used to 
compute a revised tone correction factor for that particular one-third 
octave band. Other methods of rejecting spurious tone corrections may be 
approved.
    A36.4.3.2 The tone correction procedure will underestimate EPNL if 
an important tone is of a frequency such that it is recorded in two 
adjacent one-third octave bands. An applicant must demonstrate that 
either:
    (a) No important tones are recorded in two adjacent one-third octave 
bands; or
    (b) That if an important tone has occurred, the tone correction has 
been adjusted to the value it would have had if the tone had been 
recorded fully in a single one-third octave band.
    A36.4.4 Maximum tone-corrected perceived noise level
    A36.4.4.1 The maximum tone-corrected perceived noise level, PNLTM, 
must be the maximum calculated value of the tone-corrected perceived 
noise level PNLT(k). It must be calculated using the procedure of 
section A36.4.3. To obtain a satisfactory noise time history, 
measurements must be made at 0.5 second time intervals.

    Note 1: Figure A36-2 is an example of a flyover noise time history 
where the maximum value is clearly indicated.
    Note 2: In the absence of a tone correction factor, PNLTM would 
equal PNLM.
[GRAPHIC] [TIFF OMITTED] TR08JY02.004

    A36.4.4.2 After the value of PNLTM is obtained, the frequency band 
for the largest tone correction factor is identified for the two 
preceding and two succeeding 500 ms data samples. This is performed in 
order to identity the possibility of tone suppression at PNLTM by one-
third octave band sharing of that tone. If the value of the tone 
correction factor C(k) for PNLTM is less than the average value of C(k) 
for the five consecutive time intervals, the average value of C(k) must 
be used to compute a new value for PNLTM.
    A36.4.5 Duration correction.
    A36.4.5.1 The duration correction factor D determined by the 
integration technique is defined by the expression:

[[Page 960]]

[GRAPHIC] [TIFF OMITTED] TR08JY02.005

where T is a normalizing time constant, PNLTM is the maximum value of 
          PNLT, t(1) is the first point of time after which PNLT becomes 
          greater than PNLTM-10, and t(2) is the point of time after 
          which PNLT remains constantly less than PNLTM-10.
    A36.4.5.2 Since PNLT is calculated from measured values of sound 
pressure level (SPL), there is no obvious equation for PNLT as a 
function of time. Consequently, the equation is to be rewritten with a 
summation sign instead of an integral sign as follows:
[GRAPHIC] [TIFF OMITTED] TR08JY02.006

where [Delta]t is the length of the equal increments of time for which 
          PNLT(k) is calculated and d is the time interval to the 
          nearest 0.5s during which PNLT(k) remains greater or equal to 
          PNLTM-10.
    A36.4.5.3 To obtain a satisfactory history of the perceived noise 
level use one of the following:
    (a) Half-Second time intervals for [Delta]t; or
    (b) A shorter time interval with approved limits and constants.
    A36.4.5.4 The following values for T and [Delta]t must be used in 
calculating D in the equation given in section A36.4.5.2:

T=10 s, and
[Delta]t=0.5s (or the approved sampling time interval).

Using these values, the equation for D becomes:
[GRAPHIC] [TIFF OMITTED] TR08JY02.007

where d is the duration time defined by the points corresponding to the 
          values PNLTM-10.
    A36.4.5.5 If in using the procedures given in section A36.4.5.2, the 
limits of PNLTM-10 fall between the calculated PNLT(k) values (the usual 
case), the PNLT(k) values defining the limits of the duration interval 
must be chosen from the PNLT(k) values closest to PNLTM-10. For those 
cases with more than one peak value of PNLT(k), the applicable limits 
must be chosen to yield the largest possible value for the duration 
time.
    A36.4.6 Effective perceived noise level.
    The total subjective effect of an airplane noise event, designated 
effective perceived noise level, EPNL, is equal to the algebraic sum of 
the maximum value of the tone-corrected perceived noise level, PNLTM, 
and the duration correction D. That is:

EPNL=PNLTM+D

where PNLTM and D are calculated using the procedures given in sections 
          A36.4.2, A36.4.3, A36.4.4. and A36.4.5.
    A36.4.7 Mathematical formulation of noy tables.
    A36.4.7.1 The relationship between sound pressure level (SPL) and 
the logarithm of perceived noisiness is illustrated in Figure A36-3 and 
Table A36-3.
    A36.4.7.2 The bases of the mathematical formulation are:
    (a) The slopes (M(b), M(c), M(d) and M(e)) of the straight lines;
    (b) The intercepts (SPL(b) and SPL(c)) of the lines on the SPL axis; 
and
    (c) The coordinates of the discontinuities, SPL(a) and log n(a); 
SPL(d) and log n=-1.0; and SPL(e) and log n=log (0.3).
    A36.4.7.3 Calculate noy values using the following equations:
    (a)

SPL =SPL (a)
n=antilog {(c)[SPL-SPL(c)]{time} 


[[Page 961]]


    (b)

SPL(b) <=SPL 1.5 EPNdB. No test result may be omitted from the 
averaging process unless approved by the FAA.

    Note: Permitted methods for calculating the 90 percent confidence 
interval are shown in the current advisory circular for this part.

    A36.5.4.3 The average EPNL figures obtained by the process described 
in section A36.5.4.1 must be those by which the noise performance of the 
airplane is assessed against the noise certification criteria.

              Section A36.6 Nomenclature: Symbols and Units

------------------------------------------------------------------------
        Symbol                   Unit                    Meaning
------------------------------------------------------------------------
antilog...............  ......................  Antilogarithm to the
                                                 base 10.
C(k)..................  dB....................  Tone correction factor.
                                                 The factor to be added
                                                 to PNL(k) to account
                                                 for the presence of
                                                 spectral irregularities
                                                 such as tones at the k-
                                                 th increment of time.
d.....................  s.....................  Duration time. The time
                                                 interval between the
                                                 limits of t(1) and t(2)
                                                 to the nearest 0.5
                                                 second.
D.....................  dB....................  Duration correction. The
                                                 factor to be added to
                                                 PNLTM to account for
                                                 the duration of the
                                                 noise.
EPNL..................  EPNdB.................  Effective perceived
                                                 noise level. The value
                                                 of PNL adjusted for
                                                 both spectral
                                                 irregularities and
                                                 duration of the noise.
                                                 (The unit EPNdB is used
                                                 instead of the unit
                                                 dB).
EPNLr.................  EPNdB.................  Effective perceived
                                                 noise level adjusted
                                                 for reference
                                                 conditions.
f(i)..................  Hz....................  Frequency. The
                                                 geometrical mean
                                                 frequency for the i-th
                                                 one-third octave band.

[[Page 964]]

 
F (i, k)..............  dB....................  Delta-dB. The difference
                                                 between the original
                                                 sound pressure level
                                                 and the final
                                                 background sound
                                                 pressure level in the i-
                                                 th one-third octave
                                                 band at the k-th
                                                 interval of time. In
                                                 this case, background
                                                 sound pressure level
                                                 means the broadband
                                                 noise level that would
                                                 be present in the one-
                                                 third octave band in
                                                 the absence of the
                                                 tone.
h.....................  dB....................  dB-down. The value to be
                                                 subtracted from PNLTM
                                                 that defines the
                                                 duration of the noise.
H.....................  Percent...............  Relative humidity. The
                                                 ambient atmospheric
                                                 relative humidity.
i.....................  ......................  Frequency band index.
                                                 The numerical indicator
                                                 that denotes any one of
                                                 the 24 one-third octave
                                                 bands with geometrical
                                                 mean frequencies from
                                                 50 to 10,000 Hz.
k.....................  ......................  Time increment index.
                                                 The numerical indicator
                                                 that denotes the number
                                                 of equal time
                                                 increments that have
                                                 elapsed from a
                                                 reference zero.
Log...................  ......................  Logarithm to the base
                                                 10.
log n(a)..............  ......................  Noy discontinuity
                                                 coordinate. The log n
                                                 value of the
                                                 intersection point of
                                                 the straight lines
                                                 representing the
                                                 variation of SPL with
                                                 log n.
M(b), M(c), etc.......  ......................  Noy inverse slope. The
                                                 reciprocals of the
                                                 slopes of straight
                                                 lines representing the
                                                 variation of SPL with
                                                 log n.
n.....................  noy...................  The perceived noisiness
                                                 at any instant of time
                                                 that occurs in a
                                                 specified frequency
                                                 range.
n(i,k)................  noy...................  The perceived noisiness
                                                 at the k-th instant of
                                                 time that occurs in the
                                                 i-th one-third octave
                                                 band.
n(k)..................  noy...................  Maximum perceived
                                                 noisiness. The maximum
                                                 value of all of the 24
                                                 values of n(i) that
                                                 occurs at the k-th
                                                 instant of time.
N(k)..................  noy...................  Total perceived
                                                 noisiness. The total
                                                 perceived noisiness at
                                                 the k-th instant of
                                                 time calculated from
                                                 the 24-instantaneous
                                                 values of n (i, k).
p(b), p(c), etc.......  ......................  Noy slope. The slopes of
                                                 straight lines
                                                 representing the
                                                 variation of SPL with
                                                 log n.
PNL...................  PNdB..................  The perceived noise
                                                 level at any instant of
                                                 time. (The unit PNdB is
                                                 used instead of the
                                                 unit dB).
PNL(k)................  PNdB..................  The perceived noise
                                                 level calculated from
                                                 the 24 values of SPL
                                                 (i, k), at the k-th
                                                 increment of time. (The
                                                 unit PNdB is used
                                                 instead of the unit
                                                 dB).
PNLM..................  PNdB..................  Maximum perceived noise
                                                 level. The maximum
                                                 value of PNL(k). (The
                                                 unit PNdB is used
                                                 instead of the unit
                                                 dB).
PNLT..................  TPNdB.................  Tone-corrected perceived
                                                 noise level. The value
                                                 of PNL adjusted for the
                                                 spectral irregularities
                                                 that occur at any
                                                 instant of time. (The
                                                 unit TPNdB is used
                                                 instead of the unit
                                                 dB).
PNLT(k)...............  TPNdB.................  The tone-corrected
                                                 perceived noise level
                                                 that occurs at the k-th
                                                 increment of time.
                                                 PNLT(k) is obtained by
                                                 adjusting the value of
                                                 PNL(k) for the spectral
                                                 irregularities that
                                                 occur at the k-th
                                                 increment of time. (The
                                                 unit TPNdB is used
                                                 instead of the unit
                                                 dB).
PNLTM.................  TPNdB.................  Maximum tone-corrected
                                                 perceived noise level.
                                                 The maximum value of
                                                 PNLT(k). (The unit
                                                 TPNdB is used instead
                                                 of the unit dB).
PNLTr.................  TPNdB.................  Tone-corrected perceived
                                                 noise level adjusted
                                                 for reference
                                                 conditions.
s (i, k)..............  dB....................  Slope of sound pressure
                                                 level. The change in
                                                 level between adjacent
                                                 one-third octave band
                                                 sound pressure levels
                                                 at the i-th band for
                                                 the k-th instant of
                                                 time.
[Delta]s (i, k).......  dB....................  Change in slope of sound
                                                 pressure level.
s' (i, k).............  dB....................  Adjusted slope of sound
                                                 pressure level. The
                                                 change in level between
                                                 adjacent adjusted one-
                                                 third octave band sound
                                                 pressure levels at the
                                                 i-th band for the k-th
                                                 instant of time.
s (i, k)..............  dB....................  Average slope of sound
                                                 pressure level.
SPL...................  dB re.................  Sound pressure level.
                        20 [micro]Pa..........   The sound pressure
                                                 level that occurs in a
                                                 specified frequency
                                                 range at any instant of
                                                 time.
SPL(a)................  dB re.................  Noy discontinuity
                        20 [micro]Pa..........   coordinate. The SPL
                                                 value of the
                                                 intersection point of
                                                 the straight lines
                                                 representing the
                                                 variation of SPL with
                                                 log n.
SPL(b)................  dB re.................  Noy intercept. The
SPL (c)...............  20 [micro]Pa..........   intercepts on the SPL-
                                                 axis of the straight
                                                 lines representing the
                                                 variation of SPL with
                                                 log n.
SPL (i, k)............  dB re.................  The sound pressure level
                        20 [micro]Pa..........   at the k-th instant of
                                                 time that occurs in the
                                                 i-th one-third octave
                                                 band.

[[Page 965]]

 
SPL' (i, k)...........  dB re.................  Adjusted sound pressure
                        20 [micro]Pa..........   level. The first
                                                 approximation to
                                                 background sound
                                                 pressure level in the i-
                                                 th one-third octave
                                                 band for the k-th
                                                 instant of time.
SPL(i)................  dB re.................  Maximum sound pressure
                        20 [micro]Pa..........   level. The sound
                                                 pressure level that
                                                 occurs in the i-th one-
                                                 third octave band of
                                                 the spectrum for PNLTM.
SPL(i)r...............  dB re.................  Corrected maximum sound
                        20 [micro]Pa..........   pressure level. The
                                                 sound pressure level
                                                 that occurs in the i-th
                                                 one-third octave band
                                                 of the spectrum for
                                                 PNLTM corrected for
                                                 atmospheric sound
                                                 absorption.
SPL' (i, k)...........  dB re.................  Final background sound
                        20 [micro]Pa..........   pressure level. The
                                                 second and final
                                                 approximation to
                                                 background sound
                                                 pressure level in the i-
                                                 th one-third octave
                                                 band for the k-th
                                                 instant of time.
t.....................  s.....................  Elapsed time. The length
                                                 of time measured from a
                                                 reference zero.
t(1), t(2)............  s.....................  Time limit. The
                                                 beginning and end,
                                                 respectively, of the
                                                 noise time history
                                                 defined by h.
[Delta]t..............  s.....................  Time increment. The
                                                 equal increments of
                                                 time for which PNL(k)
                                                 and PNLT(k) are
                                                 calculated.
T.....................  s.....................  Normalizing time
                                                 constant. The length of
                                                 time used as a
                                                 reference in the
                                                 integration method for
                                                 computing duration
                                                 corrections, where
                                                 T=10s.
t([deg]F) ([deg]C)....  [deg]F, [deg]C........  Temperature. The ambient
                                                 air temperature.
[alpha](i)............  dB/1000ft db/100m.....  Test atmospheric
                                                 absorption. The
                                                 atmospheric attenuation
                                                 of sound that occurs in
                                                 the i-th one-third
                                                 octave band at the
                                                 measured air
                                                 temperature and
                                                 relative humidity.
[alpha](i)o...........  dB/1000ft db/100m.....  Reference atmospheric
                                                 absorption. The
                                                 atmospheric attenuation
                                                 of sound that occurs in
                                                 the i-th one-third
                                                 octave band at a
                                                 reference air
                                                 temperature and
                                                 relative humidity.
A1....................  Degrees...............  First constant climb
                                                 angle (Gear up, speed
                                                 of at least V2+10 kt
                                                 (V2+19 km/h), takeoff
                                                 thrust).
A2....................  Degrees...............  Second constant climb
                                                 angle (Gear up, speed
                                                 of at least V2+10 kt
                                                 (V2+19 km/h), after cut-
                                                 back).
[delta]...............  Degrees...............  Thrust cutback angles.
[epsi]................                           The angles defining the
                                                 points on the takeoff
                                                 flight path at which
                                                 thrust reduction is
                                                 started and ended
                                                 respectively.
[eta].................  Degrees...............  Approach angle.
[eta]r................  Degrees...............  Reference approach
                                                 angle.
[thetas]..............  Degrees...............  Noise angle (relative to
                                                 flight path). The angle
                                                 between the flight path
                                                 and noise path. It is
                                                 identical for both
                                                 measured and corrected
                                                 flight paths.
[psi].................  Degrees...............  Noise angle (relative to
                                                 ground). The angle
                                                 between the noise path
                                                 and the ground. It is
                                                 identical for both
                                                 measured and corrected
                                                 flight paths.
[mu]..................  ......................  Engine noise emission
                                                 parameter.
[mu]r.................  ......................  Reference engine noise
                                                 emission parameter.
[Delta]1..............  EPNdB.................  PNLT correction. The
                                                 correction to be added
                                                 to the EPNL calculated
                                                 from measured data to
                                                 account for noise level
                                                 changes due to
                                                 differences in
                                                 atmospheric absorption
                                                 and noise path length
                                                 between reference and
                                                 test conditions.
[Delta]2..............  EPNdB.................  Adjustment to duration
                                                 correction. The
                                                 adjustment to be made
                                                 to the EPNL calculated
                                                 from measured data to
                                                 account for noise level
                                                 changes due to the
                                                 noise duration between
                                                 reference and test
                                                 conditions.
[Delta]3..............  EPNdB.................  Source noise adjustment.
                                                 The adjustment to be
                                                 made to the EPNL
                                                 calculated from
                                                 measured data to
                                                 account for noise level
                                                 changes due to
                                                 differences between
                                                 reference and test
                                                 engine operating
                                                 conditions.
------------------------------------------------------------------------

                 Section A36.7 Sound Attenuation in Air

    A36.7.1 The atmospheric attenuation of sound must be determined in 
accordance with the procedure presented in section A36.7.2.
    A36.7.2 The relationship between sound attenuation, frequency, 
temperature, and humidity is expressed by the following equations.
    A36.7.2(a) For calculations using the English System of Units:
    [GRAPHIC] [TIFF OMITTED] TR08JY02.010
    
and
[GRAPHIC] [TIFF OMITTED] TR08JY02.011


[[Page 966]]


where

[eta]([delta]) is listed in Table A36-4 and f0 in Table A36-
          5;
[alpha](i) is the attenuation coefficient in dB/1000 ft;
[thetas] is the temperature in [deg]F; and
H is the relative humidity, expressed as a percentage.

    A36.7.2(b) For calculations using the International System of Units 
(SI):
[GRAPHIC] [TIFF OMITTED] TR08JY02.012

and
[GRAPHIC] [TIFF OMITTED] TR08JY02.013

where

[eta]([delta]) is listed in Table A36-4 and f0 in Table A36-
          5;
[alpha](i) is the attenuation coefficient in dB/100 m;
[thetas] is the temperature in [deg]C; and
H is the relative humidity, expressed as a percentage.

    A36.7.3 The values listed in table A36-4 are to be used when 
calculating the equations listed in section A36.7.2. A term of quadratic 
interpolation is to be used where necessary.

                        Section A36.8 [Reserved]

[[Page 967]]

[GRAPHIC] [TIFF OMITTED] TR08JY02.014

        Section A36.9 Adjustment of Airplane Flight Test Results.

    A36.9.1 When certification test conditions are not identical to 
reference conditions, appropriate adjustments must be made to the 
measured noise data using the methods described in this section.
    A36.9.1.1 Adjustments to the measured noise values must be made 
using one of the methods described in sections A36.9.3 and A36.9.4 for 
differences in the following:
    (a) Attenuation of the noise along its path as affected by ``inverse 
square'' and atmospheric attenuation
    (b) Duration of the noise as affected by the distance and the speed 
of the airplane relative to the measuring point

[[Page 968]]

    (c) Source noise emitted by the engine as affected by the 
differences between test and reference engine operating conditions
    (d) Airplane/engine source noise as affected by differences between 
test and reference airspeeds. In addition to the effect on duration, the 
effects of airspeed on component noise sources must be accounted for as 
follows: for conventional airplane configurations, when differences 
between test and reference airspeeds exceed 15 knots (28 km/h) true 
airspeed, test data and/or analysis approved by the FAA must be used to 
quantify the effects of the airspeed adjustment on resulting 
certification noise levels.
    A36.9.1.2 The ``integrated'' method of adjustment, described in 
section A36.9.4, must be used on takeoff or approach under the following 
conditions:
    (a) When the amount of the adjustment (using the ``simplified'' 
method) is greater than 8 dB on flyover, or 4 dB on approach; or
    (b) When the resulting final EPNL value on flyover or approach 
(using the simplified method) is within 1 dB of the limiting noise 
levels as prescribed in section B36.5 of this part.
    A36.9.2 Flight profiles.
    As described below, flight profiles for both test and reference 
conditions are defined by their geometry relative to the ground, 
together with the associated airplane speed relative to the ground, and 
the associated engine control parameter(s) used for determining the 
noise emission of the airplane.
    A36.9.2.1 Takeoff Profile.

    Note: Figure A36-4 illustrates a typical takeoff profile.

    (a) The airplane begins the takeoff roll at point A, lifts off at 
point B and begins its first climb at a constant angle at point C. Where 
thrust or power (as appropriate) cut-back is used, it is started at 
point D and completed at point E. From here, the airplane begins a 
second climb at a constant angle up to point F, the end of the noise 
certification takeoff flight path.
    (b) Position K1 is the takeoff noise measuring station 
and AK1 is the distance from start of roll to the flyover 
measuring point. Position K2 is the lateral noise measuring 
station, which is located on a line parallel to, and the specified 
distance from, the runway center line where the noise level during 
takeoff is greatest.
    (c) The distance AF is the distance over which the airplane position 
is measured and synchronized with the noise measurements, as required by 
section A36.2.3.2 of this part.
    A36.9.2.2 Approach Profile.

    Note: Figure A36-5 illustrates a typical approach profile.

    (a) The airplane begins its noise certification approach flight path 
at point G and touches down on the runway at point J, at a distance OJ 
from the runway threshold.
    (b) Position K3 is the approach noise measuring station 
and K3O is the distance from the approach noise measurement 
point to the runway threshold.
    (c) The distance GI is the distance over which the airplane position 
is measured and synchronized with the noise measurements, as required by 
section A36.2.3.2 of this part.

[[Page 969]]

[GRAPHIC] [TIFF OMITTED] TR08JY02.015

The airplane reference point for approach measurements is the instrument 
landing system (ILS) antenna. If no ILS antenna is installed an 
alternative reference point must be approved by the FAA.
    A36.9.3 Simplified method of adjustment.
    A36.9.3.1 General. As described below, the simplified adjustment 
method consists of applying adjustments (to the EPNL, which is 
calculated from the measured data) for the differences between measured 
and reference conditions at the moment of PNLTM.
    A36.9.3.2 Adjustments to PNL and PNLT.
    (a) The portions of the test flight path and the reference flight 
path described below, and illustrated in Figure A36-6, include the noise 
time history that is relevant to the calculation of flyover and approach 
EPNL. In figure A36-6:

[[Page 970]]

    (1) XY represents the portion of the measured flight path that 
includes the noise time history relevant to the calculation of flyover 
and approach EPNL; XrYr represents the 
corresponding portion of the reference flight path.
    (2) Q represents the airplane's position on the measured flight path 
at which the noise was emitted and observed as PNLTM at the noise 
measuring station K. Qr is the corresponding position on the 
reference flight path, and Kr the reference measuring 
station. QK and QrKr are, respectively, the 
measured
[GRAPHIC] [TIFF OMITTED] TR08JY02.016

and reference noise propagation paths, Qr being determined 
from the assumption that QK and QrKr form the same 
angle [thetas] with their respective flight paths.
    (b) The portions of the test flight path and the reference flight 
path described in paragraph (b)(1) and (2), and illustrated in Figure 
A36-7(a) and (b), include the noise time history that is relevant to the 
calculation of lateral EPNL.
    (1) In figure A36-7(a), XY represents the portion of the measured 
flight path that includes the noise time history that is relevant to the 
calculation of lateral EPNL; in figure A36-7(b), 
XrYr represents the corresponding portion of the 
reference flight path.
    (2) Q represents the airplane position on the measured flight path 
at which the noise was emitted and observed as PNLTM at the noise 
measuring station K. Qr is the corresponding position on the 
reference flight path, and Kr the reference measuring 
station. QK and QrKr are, respectively, the 
measured and reference noise propagation paths. In this case 
Kr is only specified as being on a particular Lateral line; 
Kr and Qr are therefore determined from the 
assumptions that QK and QrKr:
    (i) Form the same angle [thetas] with their respective flight paths; 
and
    (ii) Form the same angle [psi] with the ground.

    Note: For the lateral noise measurement, sound propagation is 
affected not only by inverse square and atmospheric attenuation, but 
also by ground absorption and reflection effects which depend mainly on 
the angle [psi].

[[Page 971]]

[GRAPHIC] [TIFF OMITTED] TR08JY02.017

    A36.9.3.2.1 The one-third octave band levels SPL(i) comprising PNL 
(the PNL at the moment of PNLTM observed at K) must be adjusted to 
reference levels SPL(i)r as follows:
    A36.9.3.2.1(a) For calculations using the English System of Units:

SPL(i)r=SPL(i)+0.001[[alpha](i)-[alpha](i)0]QK
+0.001[alpha](i)0(QK-QrKr)
+20log(QK/QrKr)

    In this expression,
    (1) The term 0.001[[alpha](i)-[alpha](i)0]QK is the 
adjustment for the effect of the change in sound attenuation 
coefficient, and [alpha](i) and [alpha](i)0 are the 
coefficients for the test and reference atmospheric conditions 
respectively, determined under section A36.7 of this appendix;
    (2) The term 0.001[alpha](i)0(QK - 
QrKr) is the adjustment for the effect of the 
change in the noise path length on the sound attenuation;
    (3) The term 20 log(QK/QrKr) is the adjustment 
for the effect of the change in the noise path length due to the 
``inverse square'' law;
    (4) QK and QrKr are measured in feet and 
[alpha](i) and [alpha](i)0 are expressed in dB/1000 ft.
    A36.9.3.2.1(b) For calculations using the International System of 
Units:

SPL(i)r=SPL(i)+0.01[[alpha](i)-[alpha](i)0]QK
+0.01[alpha](i)0 (QK - QrKr)
+20 log(QK/QrKr)

In this expression,

[[Page 972]]

    (1) The term 0.01[[alpha](i) - [alpha](i)0]QK is the 
adjustment for the effect of the change in sound attenuation 
coefficient, and [alpha](i) and [alpha](i)0 are the 
coefficients for the test and reference atmospheric conditions 
respectively, determined under section A36.7 of this appendix;
    (2) The term 0.01[alpha](i)0(QK - 
QrKr) is the adjustment for the effect of the 
change in the noise path length on the sound attenuation;
    (3) The term 20 log(QK/QrKr) is the adjustment 
for the effect of the change in the noise path length due to the inverse 
square law;
    (4) QK and QrKr are measured in meters and 
[alpha](i) and [alpha](i)0 are expressed in dB/100 m.
    A36.9.3.2.1.1 PNLT Correction.
    (a) Convert the corrected values, SPL(i)r, to 
PNLTr;
    (b) Calculate the correction term [Delta]1 using the 
following equation:

[Delta]1=PNLTr - PNLTM

    A36.9.3.2.1.2 Add [Delta]1 arithmetically to the EPNL 
calculated from the measured data.
    A36.9.3.2.2 If, during a test flight, several peak values of PNLT 
that are within 2 dB of PNLTM are observed, the procedure defined in 
section A36.9.3.2.1 must be applied at each peak, and the adjustment 
term, calculated according to section A36.9.3.2.1, must be added to each 
peak to give corresponding adjusted peak values of PNLT. If these peak 
values exceed the value at the moment of PNLTM, the maximum value of 
such exceedance must be added as a further adjustment to the EPNL 
calculated from the measured data.
    A36.9.3.3 Adjustments to duration correction.
    A36.9.3.3.1 Whenever the measured flight paths and/or the ground 
velocities of the test conditions differ from the reference flight paths 
and/or the ground velocities of the reference conditions, duration 
adjustments must be applied to the EPNL values calculated from the 
measured data. The adjustments must be calculated as described below.
    A36.9.3.3.2 For the flight path shown in Figure A36-6, the 
adjustment term is calculated as follows:

[Delta]2=-7.5 log(QK/QrKr)+10 log(V/
          Vr)

    (a) Add [Delta]2 arithmetically to the EPNL calculated 
from the measured data.
    A36.9.3.4 Source noise adjustments.
    A36.9.3.4.1 To account for differences between the parameters 
affecting engine noise as measured in the certification flight tests, 
and those calculated or specified in the reference conditions, the 
source noise adjustment must be calculated and applied. The adjustment 
is determined from the manufacturer's data approved by the FAA. Typical 
data used for this adjustment are illustrated in Figure A36-8 that shows 
a curve of EPNL versus the engine control parameter [mu], with the EPNL 
data being corrected to all the other relevant reference conditions 
(airplane mass, speed and altitude, air temperature) and for the 
difference in noise between the test engine and the average engine (as 
defined in section B36.7(b)(7)). A sufficient number of data points over 
a range of values of [mu]r is required to calculate the 
source noise adjustments for lateral, flyover and approach noise 
measurements.
[GRAPHIC] [TIFF OMITTED] TR08JY02.018

    A36.9.3.4.2 Calculate adjustment term [Delta]3 by 
subtracting the EPNL value corresponding to the parameter [mu] from the 
EPNL value corresponding to the parameter [mu]r. Add 
[Delta]3

[[Page 973]]

arithmetically to the EPNL value calculated from the measured data.
    A36.9.3.5 Symmetry adjustments.
    A36.9.3.5.1 A symmetry adjustment to each lateral noise value 
(determined at the section B36.4(b) measurement points), is to be made 
as follows:
    (a) If the symmetrical measurement point is opposite the point where 
the highest noise level is obtained on the main lateral measurement 
line, the certification noise level is the arithmetic mean of the noise 
levels measured at these two points (see Figure A36-9(a));
    (b) If the condition described in paragraph (a) of this section is 
not met, then it is assumed that the variation of noise with the 
altitude of the airplane is the same on both sides; there is a constant 
difference between the lines of noise versus altitude on both sides (see 
figure A36-9(b)). The certification noise level is the maximum value of 
the mean between these lines.
[GRAPHIC] [TIFF OMITTED] TR08JY02.019

    A36.9.4 Integrated method of adjustment
    A36.9.4.1 General. As described in this section, the integrated 
adjustment method consists of recomputing under reference conditions 
points on the PNLT time history corresponding to measured points 
obtained during the tests, and computing EPNL directly for the new time 
history obtained in this way. The main principles are described in 
sections A36.9.4.2 through A36.9.4.4.1.
    A36.9.4.2 PNLT computations.
    (a) The portions of the test flight path and the reference flight 
path described in paragraph (a)(1) and (2), and illustrated in Figure 
A36-10, include the noise time history that is relevant to the 
calculation of flyover and approach EPNL. In figure A36-10:

[[Page 974]]

[GRAPHIC] [TIFF OMITTED] TR08JY02.020

    (1) XY represents the portion of the measured flight path that 
includes the noise time history relevant to the calculation of flyover 
and approach EPNL; XrYr represents the 
corresponding reference flight path.
    (2) The points Q0, Q1, Qn represent 
airplane positions on the measured flight path at time t0, 
t1 and tn respectively. Point Q1 is the 
point at which the noise was emitted and observed as one-third octave 
values SPL(i)1 at the noise measuring station K at time 
t1. Point Qr1 represents the corresponding 
position on the reference flight path for noise observed as 
SPL(i)r1 at the reference measuring station Kr at 
time tr1. Q1K and Qr1Kr are 
respectively the measured and reference noise propagation paths, which 
in each case form the angle [thetas]1 with their respective 
flight paths. Qr0 and Qrn are similarly the points 
on the reference flight path corresponding to Q0 and 
Qn on the measured flight path. Q0 and 
Qn are chosen so that between Qr0 and 
Qrn all values of PNLTr (computed as described in 
paragraphs A36.9.4.2.2 and A36.9.4.2.3) within 10 dB of the peak value 
are included.
    (b) The portions of the test flight path and the reference flight 
path described in paragraph (b)(1) and (2), and illustrated in Figure 
A36-11(a) and (b), include the noise time history that is relevant to 
the calculation of lateral EPNL.
    (1) In figure A36-11(a) XY represents the portion of the measured 
flight path that includes the noise time history that is relevant to the 
calculation of lateral EPNL; in figure A36-11(b), 
XrYr represents the corresponding portion of the 
reference flight path.
    (2) The points Q0, Q1 and Qn 
represent airplane positions on the measured flight path at time 
t0, t1 and tn respectively. Point 
Q1 is the point at which the noise was emitted and observed 
as one-third octave values SPL(i)1 at the noise measuring 
station K at time t1. The point Qr1 represents the 
corresponding position on the reference flight path for noise observed 
as SPL(i)r1 at the measuring station Kr at time 
tr1. Q1K and Qr1Kr are 
respectively the measured and reference noise propagation paths. 
Qr0 and Qrn are similarly the points on the 
reference flight path corresponding to Q0 and Qn 
on the measured flight path.

[[Page 975]]

[GRAPHIC] [TIFF OMITTED] TR08JY02.021

    Q0 and Qn are chosen to that between 
Qro and Qrn all values of PNLTr 
(computed as described in paragraphs A36.9.4.2.2 and A36.9.4.2.3) within 
10 dB of the peak value are included. In this case Kr is only 
specified as

[[Page 976]]

being on a particular lateral line. The position of Kr and 
Qr1 are determined from the following requirements.
    (i) Q1K and Qr1Kr form the same 
angle [thetas]1 with their respective flight paths; and
    (ii) The differences between the angles 1 and 
r1 must be minimized using a method, approved by the FAA. The 
differences between the angles are minimized since, for geometrical 
reasons, it is generally not possible to choose Kr so that 
the condition described in paragraph A36.9.4.2(b)(2)(i) is met while at 
the same time keeping 1 and r1 equal.

    Note: For the lateral noise measurement, sound propagation is 
affected not only by ``inverse square'' and atmospheric attenuation, but 
also by ground absorption and reflection effects which depend mainly on 
the angle.

    A36.9.4.2.1 In paragraphs A36.9.4.2(a)(2) and (b)(2) the time 
tr1 is later (for Qr1Kr 
Q1K) than t1 by two separate amounts:
    (1) The time taken for the airplane to travel the distance 
Qr1Qr0 at a speed Vr less the time 
taken for it to travel Q1Q0 at V;
    (2) The time taken for sound to travel the distance 
Qr1Kr-Q1K.

    Note: For the flight paths described in paragraphs A36.9.4.2(a) and 
(b), the use of thrust or power cut-back will result in test and 
reference flight paths at full thrust or power and at cut-back thrust or 
power. Where the transient region between these thrust or power levels 
affects the final result, an interpolation must be made between them by 
an approved method such as that given in the current advisory circular 
for this part.

    A36.9.4.2.2 The measured values of SPL(i)1 must be 
adjusted to the reference values SPL(i)r1 to account for the 
differences between measured and reference noise path lengths and 
between measured and reference atmospheric conditions, using the methods 
of section A36.9.3.2.1 of this appendix. A corresponding value of 
PNLr1 must be computed according to the method in section 
A36.4.2. Values of PNLr must be computed for times 
t0 through tn.
    A36.9.4.2.3 For each value of PNLr1, a tone correction 
factor C1 must be determined by analyzing the reference 
values SPL(i)r using the methods of section A36.4.3 of this 
appendix, and added to PNLr1 to yield PNLTr1. 
Using the process described in this paragraph, values of 
PNLTr must be computed for times t0 through 
tn.
    A36.9.4.3 Duration correction.
    A36.9.4.3.1 The values of PNLTr corresponding to those of 
PNLT at each one-half second interval must be plotted against time 
(PNLTr1 at time tr1). The duration correction must 
then be determined using the method of section A36.4.5.1 of this 
appendix, to yield EPNLr.
    A36.9.4.4 Source Noise Adjustment.
    A36.9.4.4.1 A source noise adjustment, [Delta]3, must be 
determined using the methods of section A36.9.3.4 of this appendix.

              A36.9.5 Flight Path Identification Positions
------------------------------------------------------------------------
            Position                            Description
------------------------------------------------------------------------
A...............................  Start of Takeoff roll.
B...............................  Lift-off.
C...............................  Start of first constant climb.
D...............................  Start of thrust reduction.
E...............................  Start of second constant climb.
F...............................  End of noise certification Takeoff
                                   flight path.
G...............................  Start of noise certification Approach
                                   flight path.
H...............................  Position on Approach path directly
                                   above noise measuring station.
I...............................  Start of level-off.
J...............................  Touchdown.
K...............................  Noise measurement point.
Kr..............................  Reference measurement point.
K1..............................  Flyover noise measurement point.
K2..............................  Lateral noise measurement point.
K3..............................  Approach noise measurement point.
M...............................  End of noise certification Takeoff
                                   flight track.
O...............................  Threshold of Approach end of runway.
P...............................  Start of noise certification Approach
                                   flight track.
Q...............................  Position on measured Takeoff flight
                                   path corresponding to apparent PNLTM
                                   at station K See section A36.9.3.2.
Qr..............................  Position on corrected Takeoff flight
                                   path corresponding to PNLTM at
                                   station K. See section A36.9.3.2.
V...............................  Airplane test speed.
Vr..............................  Airplane reference speed.
------------------------------------------------------------------------


                      A36.9.6 Flight Path Distances
------------------------------------------------------------------------
      Distance                 Unit                     Meaning
------------------------------------------------------------------------
AB.................  Feet (meters)..........  Length of takeoff roll.
                                               The distance along the
                                               runway between the start
                                               of takeoff roll and lift
                                               off.
AK.................  Feet (meters)..........  Takeoff measurement
                                               distance. The distance
                                               from the start of roll to
                                               the takeoff noise
                                               measurement station along
                                               the extended center line
                                               of the runway.
AM.................  Feet (meters)..........  Takeoff flight track
                                               distance. The distance
                                               from the start of roll to
                                               the takeoff flight track
                                               position along the
                                               extended center line of
                                               the runway after which
                                               the position of the
                                               airplane need no longer
                                               be recorded.
QK.................  Feet (meters)..........  Measured noise path. The
                                               distance from the
                                               measured airplane
                                               position Q to station K.
QrKr...............  Feet (meters)..........  Reference noise path. The
                                               distance from the
                                               reference airplane
                                               position Qr to station
                                               Kr.
K3H................  Feet (meters)..........  Airplane approach height.
                                               The height of the
                                               airplane above the
                                               approach measuring
                                               station.
OK3................  Feet (meters)..........  Approach measurement
                                               distance. The distance
                                               from the runway threshold
                                               to the approach
                                               measurement station along
                                               the extended center line
                                               of the runway.

[[Page 977]]

 
OP.................  Feet (meters)..........  Approach flight track
                                               distance. The distance
                                               from the runway threshold
                                               to the approach flight
                                               track position along the
                                               extended center line of
                                               the runway after which
                                               the position of the
                                               airplane need no longer
                                               be recorded.
------------------------------------------------------------------------


[Amdt. 36-54, 67 FR 45212, July 8, 2002; Amdt. 36-24, 67 FR 63195, 
63196, Oct. 10, 2002; 68 FR 1512, Jan. 10, 2003; Amdt. 36-26, 70 FR 
38749, July 5, 2005]



Sec. Appendix B to Part 36--Noise Levels for Transport Category and Jet 
                      Airplanes Under Sec. 36.103

Sec.

B36.1 Noise Measurement and Evaluation.
B36.2 Noise Evaluation Metric.
B36.3 Reference Noise Measurement Points.
B36.4 Test Noise Measurement Points.
B36.5 Maximum Noise Levels.
B36.6 Trade-Offs.
B36.7 Noise Certification Reference Procedures and Conditions.
B36.8 Noise Certification Test Procedures.

             Section B36.1 Noise measurement and evaluation

    (a) The procedures of Appendix A of this part, or approved 
equivalent procedures, must be used to determine noise levels of an 
airplane. These noise levels must be used to show compliance with the 
requirements of this appendix.
    (b) For Stage 4 airplanes, an acceptable alternative for noise 
measurement and evaluation is Appendix 2 to the International Civil 
Aviation Organization (ICAO) Annex 16, Environmental Protection, Volume 
I, Aircraft Noise, Third Edition, July 1993, Amendment 7, effective 
March 21, 2002. [Incorporated by reference, see Sec. 36.6].

                  Section B36.2 Noise Evaluation Metric

    The noise evaluation metric is the effective perceived noise level 
expressed in EPNdB, as calculated using the procedures of appendix A of 
this part.

            Section B36.3 Reference Noise Measurement Points

    When tested using the procedures of this part, except as provided in 
section B36.6, an airplane may not exceed the noise levels specified in 
section B36.5 at the following points on level terrain:
    (a) Lateral full-power reference noise measurement point:
    (1) For jet airplanes: The point on a line parallel to and 1,476 
feet (450 m) from the runway centerline, or extended centerline, where 
the noise level after lift-off is at a maximum during takeoff. For the 
purpose of showing compliance with Stage 1 or Stage 2 noise limits for 
an airplane powered by more than three jet engines, the distance from 
the runway centerline must be 0.35 nautical miles (648 m). For jet 
airplanes, when approved by the FAA, the maximum lateral noise at 
takeoff thrust may be assumed to occur at the point (or its approved 
equivalent) along the extended centerline of the runway where the 
airplane reaches 985 feet (300 meters) altitude above ground level. A 
height of 1427 feet (435 meters) may be assumed for Stage 1 or Stage 2 
four engine airplanes. The altitude of the airplane as it passes the 
noise measurement points must be within +328 to -164 feet (+100 to -50 
meters) of the target altitude. For airplanes powered by other than jet 
engines, the altitude for maximum lateral noise must be determined 
experimentally.
    (2) For propeller-driven airplanes: The point on the extended 
centerline of the runway above which the airplane, at full takeoff 
power, reaches a height of 2,133 feet (650 meters). For tests conducted 
before August 7, 2002, an applicant may use the measurement point 
specified in section B36.3(a)(1) as an alternative.
    (b) Flyover reference noise measurement point: The point on the 
extended centerline of the runway that is 21,325 feet (6,500 m) from the 
start of the takeoff roll;
    (c) Approach reference noise measurement point: The point on the 
extended centerline of the runway that is 6,562 feet (2,000 m) from the 
runway threshold. On level ground, this corresponds to a position that 
is 394 feet (120 m) vertically below the 3[deg] descent path, which 
originates at a point on the runway 984 feet (300 m) beyond the 
threshold.

              Section B36.4 Test noise measurement points.

    (a) If the test noise measurement points are not located at the 
reference noise measurement points, any corrections for the difference 
in position are to be made using the same adjustment procedures as for 
the differences between test and reference flight paths.
    (b) The applicant must use a sufficient number of lateral test noise 
measurement points to demonstrate to the FAA that the maximum noise 
level on the appropriate lateral line has been determined. For jet 
airplanes, simultaneous measurements must be made at one test noise 
measurement point at its symmetrical point on the other side of the 
runway. Propeller-driven airplanes have an inherent asymmetry in lateral 
noise. Therefore, simultaneous measurements must be made at each and 
every test noise measurement point at its symmetrical position on the 
opposite side of the runway. The measurement points are considered to be

[[Page 978]]

symmetrical if they are longitudinally within 33 feet (10 meters) of each other.

                   Section B36.5 Maximum Noise Levels

    Except as provided in section B36.6 of this appendix, maximum noise 
levels, when determined in accordance with the noise evaluation methods 
of appendix A of this part, may not exceed the following:
    (a) For acoustical changes to Stage 1 airplanes, regardless of the 
number of engines, the noise levels prescribed under Sec. 36.7(c) of 
this part.
    (b) For any Stage 2 airplane regardless of the number of engines:
    (1) Flyover: 108 EPNdB for maximum weight of 600,000 pounds or more; 
for each halving of maximum weight (from 600,000 pounds), reduce the 
limit by 5 EPNdB; the limit is 93 EPNdB for a maximum weight of 75,000 
pounds or less.
    (2) Lateral and approach: 108 EPNdB for maximum weight of 600,000 
pounds or more; for each halving of maximum weight (from 600,000 
pounds), reduce the limit by 2 EPNdB; the limit is 102 EPNdB for a 
maximum weight of 75,000 pounds or less.
    (c) For any Stage 3 airplane:
    (1) Flyover.
    (i) For airplanes with more than 3 engines: 106 EPNdB for maximum 
weight of 850,000 pounds or more; for each halving of maximum weight 
(from 850,000 pounds), reduce the limit by 4 EPNdB; the limit is 89 
EPNdB for a maximum weight of 44,673 pounds or less;
    (ii) For airplanes with 3 engines: 104 EPNdB for maximum weight of 
850,000 pounds or more; for each halving of maximum weight (from 850,000 
pounds), reduce the limit by 4 EPNdB; the limit is 89 EPNdB for a 
maximum weight of 63,177 pounds or less; and
    (iii) For airplanes with fewer than 3 engines: 101 EPNdB for maximum 
weight of 850,000 pounds or more; for each halving of maximum weight 
(from 850,000 pounds), reduce the limit by 4 EPNdB; the limit is 89 
EPNdB for a maximum weight of 106,250 pounds or less.
    (2) Lateral, regardless of the number of engines: 103 EPNdB for 
maximum weight of 882,000 pounds or more; for each halving of maximum 
weight (from 882,000 pounds), reduce the limit by 2.56 EPNdB; the limit 
is 94 EPNdB for a maximum weight of 77,200 pounds or less.
    (3) Approach, regardless of the number of engines: 105 EPNdB for 
maximum weight of 617,300 pounds or more; for each halving of maximum 
weight (from 617,300 pounds), reduce the limit by 2.33 EPNdB; the limit 
is 98 EPNdB for a maximum weight of 77,200 pounds or less.
    (d) For any Stage 4 airplane, the flyover, lateral, and approach 
maximum noise levels are prescribed in Chapter 4, Paragraph 4.4, Maximum 
Noise Levels, and Chapter 3, Paragraph 3.4, Maximum Noise Levels, of the 
International Civil Aviation Organization (ICAO) Annex 16, Environmental 
Protection, Volume I, Aircraft Noise, Third Edition, July 1993, 
Amendment 7, effective March 21, 2002. [Incorporated by reference, see 
Sec. 36.6].

                        Section B36.6 Trade-Offs

    Except when prohibited by sections 36.7(c)(1) and 36.7(d)(1)(ii), if 
the maximum noise levels are exceeded at any one or two measurement 
points, the following conditions must be met:
    (a) The sum of the exceedance(s) may not be greater than 3 EPNdB;
    (b) Any exceedance at any single point may not be greater than 2 
EPNdB, and
    (c) Any exceedance(s) must be offset by a corresponding amount at 
another point or points.

  Section B36.7 Noise Certification Reference Procedures and Conditions

    (a) General conditions:
    (1) All reference procedures must meet the requirements of section 
36.3 of this part.
    (2) Calculations of airplane performance and flight path must be 
made using the reference procedures and must be approved by the FAA.
    (3) Applicants must use the takeoff and approach reference 
procedures prescribed in paragraphs (b) and (c) of this section.
    (4) [Reserved]
    (5) The reference procedures must be determined for the following 
reference conditions. The reference atmosphere is homogeneous in terms 
of temperature and relative humidity when used for the calculation of 
atmospheric absorption coefficients.
    (i) Sea level atmospheric pressure of 2116 pounds per square foot 
(psf) (1013.25 hPa);
    (ii) Ambient sea-level air temperature of 77 [deg]F (25 [deg]C, 
i.e., ISA+10 [deg]C);
    (iii) Relative humidity of 70 per cent;
    (iv) Zero wind.
    (v) In defining the reference takeoff flight path(s) for the takeoff 
and lateral noise measurements, the runway gradient is zero.
    (b) Takeoff reference procedure:
    The takeoff reference flight path is to be calculated using the 
following:
    (1) Average engine takeoff thrust or power must be used from the 
start of takeoff to the point where at least the following height above 
runway level is reached. The takeoff thrust/power used must be the 
maximum available for normal operations given in the performance section 
of the airplane flight manual under the reference atmospheric conditions 
given in section B36.7(a)(5).
    (i) For Stage 1 airplanes and for Stage 2 airplanes that do not have 
jet engines with a bypass ratio of 2 or more, the following apply:

[[Page 979]]

    (A): For airplanes with more than three jet engines--700 feet (214 
meters).
    (B): For all other airplanes--1,000 feet (305 meters).
    (ii) For Stage 2 airplanes that have jet engines with a bypass ratio 
of 2 or more and for Stage 3 airplanes, the following apply:
    (A): For airplanes with more than three engines--689 feet (210 
meters).
    (B): For airplanes with three engines--853 feet (260 meters).
    (C): For airplanes with fewer than three engines--984 feet (300 
meters).
    (2) Upon reaching the height specified in paragraph (b)(1) of this 
section, airplane thrust or power must not be reduced below that 
required to maintain either of the following, whichever is greater:
    (i) A climb gradient of 4 per cent; or
    (ii) In the case of multi-engine airplanes, level flight with one 
engine inoperative.
    (3) For the purpose of determining the lateral noise level, the 
reference flight path must be calculated using full takeoff power 
throughout the test run without a reduction in thrust or power. For 
tests conducted before August 7, 2002, a single reference flight path 
that includes thrust cutback in accordance with paragraph (b)(2) of this 
section, is an acceptable alternative in determining the lateral noise 
level.
    (4) The takeoff reference speed is the all-engine operating takeoff 
climb speed selected by the applicant for use in normal operation; this 
speed must be at least V2+10kt (V2+19km/h) but may not be greater than 
V2+20kt (V2+37km/h). This speed must be attained as soon as practicable 
after lift-off and be maintained throughout the takeoff noise 
certification test. For Concorde airplanes, the test day speeds and the 
acoustic day reference speed are the minimum approved value of V2+35 
knots, or the all-engines-operating speed at 35 feet, whichever speed is 
greater as determined under the regulations constituting the type 
certification basis of the airplane; this reference speed may not exceed 
250 knots. For all airplanes, noise values measured at the test day 
speeds must be corrected to the acoustic day reference speed.
    (5) The takeoff configuration selected by the applicant must be 
maintained constantly throughout the takeoff reference procedure, except 
that the landing gear may be retracted. Configuration means the center 
of gravity position, and the status of the airplane systems that can 
affect airplane performance or noise. Examples include, the position of 
lift augmentation devices, whether the APU is operating, and whether air 
bleeds and engine power take-offs are operating;
    (6) The weight of the airplane at the brake release must be the 
maximum takeoff weight at which the noise certification is requested, 
which may result in an operating limitation as specified in Sec. 
36.1581(d); and
    (7) The average engine is defined as the average of all the 
certification compliant engines used during the airplane flight tests, 
up to and during certification, when operating within the limitations 
and according to the procedures given in the Flight Manual. This will 
determine the relationship of thrust/power to control parameters (e.g., 
N1 or EPR). Noise measurements made during certification 
tests must be corrected using this relationship.
    (c) Approach reference procedure:
    The approach reference flight path must be calculated using the 
following:
    (1) The airplane is stabilized and following a 3[deg] glide path;
    (2) For subsonic airplanes, a steady approach speed of 
Vref + 10 kts (Vref + 19 km/h) with thrust and 
power stabilized must be established and maintained over the approach 
measuring point. Vref is the reference landing speed, which 
is defined as the speed of the airplane, in a specified landing 
configuration, at the point where it descends through the landing screen 
height in the determination of the landing distance for manual landings. 
For Concorde airplanes, a steady approach speed that is either the 
landing reference speed + 10 knots or the speed used in establishing the 
approved landing distance under the airworthiness regulations 
constituting the type certification basis of the airplane, whichever 
speed is greater. This speed must be established and maintained over the 
approach measuring point.
    (3) The constant approach configuration used in the airworthiness 
certification tests, but with the landing gear down, must be maintained 
throughout the approach reference procedure;
    (4) The weight of the airplane at touchdown must be the maximum 
landing weight permitted in the approach configuration defined in 
paragraph (c)(3) of this section at which noise certification is 
requested, except as provided in Sec. 36.1581(d) of this part; and
    (5) The most critical configuration must be used; this configuration 
is defined as that which produces the highest noise level with normal 
deployment of aerodynamic control surfaces including lift and drag 
producing devices, at the weight at which certification is requested. 
This configuration includes all those items listed in section A36.5.2.5 
of appendix A of this part that contribute to the noisiest continuous 
state at the maximum landing weight in normal operation.

            Section B36.8 Noise Certification Test Procedures

    (a) All test procedures must be approved by the FAA.
    (b) The test procedures and noise measurements must be conducted and 
processed in an approved manner to yield the noise evaluation metric 
EPNL, in units of EPNdB, as described in appendix A of this part.

[[Page 980]]

    (c) Acoustic data must be adjusted to the reference conditions 
specified in this appendix using the methods described in appendix A of 
this part. Adjustments for speed and thrust must be made as described in 
section A36.9 of this part.
    (d) If the airplane's weight during the test is different from the 
weight at which noise certification is requested, the required EPNL 
adjustment may not exceed 2 EPNdB for each takeoff and 1 EPNdB for each 
approach. Data approved by the FAA must be used to determine the 
variation of EPNL with weight for both takeoff and approach test 
conditions. The necessary EPNL adjustment for variations in approach 
flight path from the reference flight path must not exceed 2 EPNdB.
    (e) For approach, a steady glide path angle of 3[deg] 0.5[deg] is acceptable.
    (f) If equivalent test procedures different from the reference 
procedures are used, the test procedures and all methods for adjusting 
the results to the reference procedures must be approved by the FAA. The 
adjustments may not exceed 16 EPNdB on takeoff and 8 EPNdB on approach. 
If the adjustment is more than 8 EPNdB on takeoff, or more than 4 EPNdB 
on approach, the resulting numbers must be more than 2 EPNdB below the 
limit noise levels specified in section B36.5.
    (g) During takeoff, lateral, and approach tests, the airplane 
variation in instantaneous indicated airspeed must be maintained within 
3% of the average airspeed between the 10 dB-down 
points. This airspeed is determined by the pilot's airspeed indicator. 
However, if the instantaneous indicated airspeed exceeds 3 kt (5.5 km/h) of the average 
airspeed over the 10 dB-down points, and is determined by the FAA 
representative on the flight deck to be due to atmospheric turbulence, 
then the flight so affected must be rejected for noise certification 
purposes.

    Note: Guidance material on the use of equivalent procedures is 
provided in the current advisory circular for this part.

[Amdt. 36-54, 67 FR 45235, July 8, 2002; Amdt. 36-24, 67 FR 63196, Oct. 
10, 2002; 68 FR 1512, Jan. 10, 2003; Amdt. 36-26, 70 FR 38749, July 5, 
2005]



                Sec. Appendixes C-E to Part 36 [Reserved]



  Sec. Appendix F to Part 36--Flyover Noise Requirements for Propeller-
 Driven Small Airplane and Propeller-Driven, Commuter Category Airplane 
             Certification Tests Prior to December 22, 1988

                             part a--general

Sec.
F36.1 Scope.

                        part b--noise measurement

F36.101 General test conditions.
F36.103 Acoustical measurement system.
F36.105 Sensing, recording, and reproducing equipment.
F36.107 Noise measurement procedures.
F36.109 Data recording, reporting, and approval.
F36.111 Flight procedures.

                         part c--data correction

F36.201 Correction of data.
F36.203 Validity of results.

                          part d--noise limits

F36.301 Aircraft noise limits.

                             part a--general

    Section F36.1 Scope. This appendix prescribes noise level limits and 
procedures for measuring and correcting noise data for the propeller 
driven small airplanes specified in Sec. Sec. 36.1 and 36.501(b).

                        part b--noise measurement

                  Sec. F36.101 General test conditions.

    (a) The test area must be relatively flat terrain having no 
excessive sound absorption characteristics such as those caused by 
thick, matted, or tall grass, by shrubs, or by wooded areas. No 
obstructions which significantly influence the sound field from the 
airplane may exist within a conical space above the measurement 
position, the cone being defined by an axis normal to the ground and by 
a half-angle 75 degrees from this axis.
    (b) The tests must be carried out under the following conditions:
    (1) There may be no precipitation.
    (2) Relative humidity may not be higher than 90 percent or lower 
than 30 percent.
    (3) Ambient temperature may not be above 86 degrees F. or below 41 
degrees F. at 33[foot] above ground. If the measurement site is within 1 
n.m. of an airport thermometer the airport reported temperature may be 
used.
    (4) Reported wind may not be above 10 knots at 33[foot] above 
ground. If wind velocities of more than 4 knots are reported, the flight 
direction must be aligned to within 15 degrees of 
wind direction and flights with tail wind and head wind must be made in 
equal numbers. If the measurement site is within 1 n.m. of an airport 
anemometer, the airport reported wind may be used.
    (5) There may be no temperature inversion or anomalous wind 
conditions that would significantly alter the noise level of the 
airplane when the noise is recorded at the required measuring point.
    (6) The flight test procedures, measuring equipment, and noise 
measurement procedures must be approved by the FAA.
    (7) Sound pressure level data for noise evaluation purposes must be 
obtained with

[[Page 981]]

acoustical equipment that complies with section F36.103 of this 
appendix.

    Sec. F36.103 Acoustical measurement system. The acoustical 
measurement system must consist of approved equipment equivalent to the 
following:
    (a) A microphone system with frequency response compatible with 
measurement and analysis system accuracy as prescribed in section 
F36.105 of this appendix.
    (b) Tripods or similar microphone mountings that minimize 
interference with the sound being measured.
    (c) Recording and reproducing equipment characteristics, frequency 
response, and dynamic range compatible with the response and accuracy 
requirements of section F36.105 of this appendix.
    (d) Acoustic calibrators using sine wave or broadband noise of known 
sound pressure level. If broadband noise is used, the signal must be 
described in terms of its average and maximum root-mean-square (rms) 
value for nonoverload signal level.

       Sec. F36.105 Sensing, recording, and reproducing equipment.

    (a) The noise produced by the airplane must be recorded. A magnetic 
tape recorder is acceptable.
    (b) The characteristics of the system must comply with the 
recommendations in International Electrotechnical Commission (IEC) 
Publication No. 179, entitled ``Precision Sound Level Meters'' as 
incorporated by reference in Part 36 under Sec. 36.6 of this part.
    (c) The response of the complete system to a sensibly plane 
progressive sinusoidal wave of constant amplitude must lie within the 
tolerance limits specified in IEC Publication No. 179, dated 1973, over 
the frequency range 45 to 11,200 Hz.
    (d) If limitations of the dynamic range of the equipment make it 
necessary, high frequency pre-emphasis must be added to the recording 
channel with the converse de-emphasis on playback. The pre-emphasis must 
be applied such that the instantaneous recorded sound pressure level of 
the noise signal between 800 and 11,200 Hz does not vary more than 20 dB 
between the maximum and minimum one-third octave bands.
    (e) If requested by the Administrator, the recorded noise signal 
must be read through an ``A'' filter with dynamic characteristics 
designated ``slow,'' as defined in IEC Publication No. 179, dated 1973. 
The output signal from the filter must be fed to a rectifying circuit 
with square law rectification, integrated with time constants for charge 
and discharge of about 1 second or 800 milliseconds.
    (f) The equipment must be acoustically calibrated using facilities 
for acoustic freefield calibration and if analysis of the tape recording 
is requested by the Administrator, the analysis equipment shall be 
electronically calibrated by a method approved by the FAA.
    (g) A windscreen must be employed with microphone during all 
measurements of aircraft noise when the wind speed is in excess of 6 
knots.

               Sec. F36.107 Noise measurement procedures.

    (a) The microphones must be oriented in a known direction so that 
the maximum sound received arrives as nearly as possible in the 
direction for which the microphones are calibrated. The microphone 
sensing elements must be approximately 4[foot] above ground.
    (b) Immediately prior to and after each test; a recorded acoustic 
calibration of the system must be made in the field with an acoustic 
calibrator for the two purposes of checking system sensitivity and 
providing an acoustic reference level for the analysis of the sound 
level data.
    (c) The ambient noise, including both acoustical background and 
electrical noise of the measurement systems, must be recorded and 
determined in the test area with the system gain set at levels that will 
be used for aircraft noise measurements. If aircraft sound pressure 
levels do not exceed the background sound pressure levels by at least 10 
dB(A), approved corrections for the contribution of background sound 
pressure level to the observed sound pressure level must be applied.

          Sec. F36.109 Data recording, reporting, and approval.

    (a) Data representing physical measurements or corrections to 
measured data must be recorded in permanent form and appended to the 
record except that corrections to measurements for normal equipment 
response deviations need not be reported. All other corrections must be 
approved. Estimates must be made of the individual errors inherent in 
each of the operations employed in obtaining the final data.
    (b) Measured and corrected sound pressure levels obtained with 
equipment conforming to the specifications described in section F36.105 
of this appendix must be reported.
    (c) The type of equipment used for measurement and analysis of all 
acoustic, airplane performance, and meteorological data must be 
reported.
    (d) The following atmospheric data, measured immediately before, 
after, or during each test at the observation points prescribed in 
section F36.101 of this appendix must be reported:
    (1) Air temperature and relative humidity.
    (2) Maximum, minimum, and average wind velocities.
    (e) Comments on local topography, ground cover, and events that 
might interfere with sound recordings must be reported.

[[Page 982]]

    (f) The following airplane information must be reported:
    (1) Type, model and serial numbers (if any) of airplanes, engines, 
and propellers.
    (2) Any modifications or nonstandard equipment likely to affect the 
noise characteristics of the airplane.
    (3) Maximum certificated takeoff weights.
    (4) Airspeed in knots for each overflight of the measuring point.
    (5) Engine performance in terms of revolutions per minute and other 
relevant parameters for each overflight.
    (6) Aircraft height in feet determined by a calibrated altimeter in 
the aircraft, approved photographic techniques, or approved tracking 
facilities.
    (g) Aircraft speed and position and engine performance parameters 
must be recorded at an approved sampling rate sufficient to ensure 
compliance with the test procedures and conditions of this appendix.

                     Sec. F36.111 Flight procedures.

    (a) Tests to demonstrate compliance with the noise level 
requirements of this appendix must include at least six level flights 
over the measuring station at a height of 1,000[foot] 30[foot] and 10 degrees from the 
zenith when passing overhead.
    (b) Each test over flight must be conducted:
    (1) At not less than the highest power in the normal operating range 
provided in an Airplane Flight Manual, or in any combination of approved 
manual material, approved placard, or approved instrument markings; and
    (2) At stabilized speed with propellers synchronized and with the 
airplane in cruise configuration, except that if the speed at the power 
setting prescribed in this paragraph would exceed the maximum speed 
authorized in level flight, accelerated flight is acceptable.

                         part c--data correction

                    Sec. F36.201 Correction of data.

    (a) Noise data obtained when the temperature is outside the range of 
68 degrees F. 9 degrees F., or the relative 
humidity is below 40 percent, must be corrected to 77 degrees F. and 70 
percent relative humidity by a method approved by the FAA.
    (b) The performance correction prescribed in paragraph (c) of this 
section must be used. It must be determined by the method described in 
this appendix, and must be added algebraically to the measured value. It 
is limited to 5dB(A).
    (c) The performance correction must be computed by using the 
following formula:
[GRAPHIC] [TIFF OMITTED] TC28SE91.113

Where:

D50=Takeoff distance to 50 feet at maximum certificated 
          takeoff weight.
R/C=Certificated best rate of climb (fpm).
Vy=Speed for best rate of climb in the same units as rate of climb.

    (d) When takeoff distance to 50[foot] is not listed as approved 
performance information, the figures of 2000 for single-engine airplanes 
and 1600[foot] for multi-engine airplanes must be used.

Sec. F36.203 Validity of results.

    (a) The test results must produce an average dB(A) and its 90 
percent confidence limits, the noise level being the arithmetic average 
of the corrected acoustical measurements for all valid test runs over 
the measuring point.
    (b) The samples must be large enough to establish statistically a 90 
pecent confidence limit not to exceed 1.5 dB(A). 
No test result may be omitted from the averaging process, unless 
omission is approved by the FAA.

                          part d--noise limits

                   Sec. F36.301 Aircraft noise limits.

    (a) Compliance with this section must be shown with noise data 
measured and corrected as prescribed in Parts B and C of this appendix.
    (b) For airplanes for which application for a type certificate is 
made on or after October 10, 1973, the noise level must not exceed 68 
dB(A) up to and including aircraft weights of 1,320 pounds (600 kg.). 
For weights greater than 1,320 pounds up to and including 3,630 pounds 
(1.650 kg.) the limit increases at the rate of 1 dB/165 pounds (1 dB/75 
kg.) to 82 dB(A) at 3,630 pounds, after which it is constant at 82 
dB(A). However, airplanes produced under type certificates covered by 
this paragraph must also meet paragraph (d) of this section for the 
original issuance of standard airworthiness certificates or restricted 
category airworthiness certificates if those airplanes have not had 
flight time before the date specified in that paragraph.

[[Page 983]]

    (c) For airplanes for which application for a type certificate is 
made on or after January 1, 1975, the noise levels may not exceed the 
noise limit curve prescribed in paragraph (b) of this section, except 
that 80 dB(A) may not be exceeded.
    (d) For airplanes for which application is made for a standard 
airworthiness certificate or for a restricted category airworthiness 
certificate, and that have not had any flight time before January 1, 
1980, the requirements of paragraph (c) of this section apply, 
regardless of date of application, to the original issuance of the 
certificate for that airplane.

[Doc. No. 13243, 40 FR 1035, Jan. 6, 1975; 40 FR 6347, Feb. 11, 1975, as 
amended by Amdt. 36-6, 41 FR 56064, Dec. 23, 1976; Amdt. 36-6, 42 FR 
4113, Jan. 24, 1977; Amdt. 36-9, 43 FR 8754, Mar. 2, 1978; Amdt. 36-13, 
52 FR 1836, Jan. 15, 1987; Amdt. 36-16, 53 FR 47400, Nov. 22, 1988]



  Sec. Appendix G to Part 36--Takeoff Noise Requirements for Propeller-
 Driven Small Airplane and Propeller-Driven, Commuter Category Airplane 
            Certification Tests on or After December 22, 1988

                             part a--general

Sec.
G36.1 Scope.

                        part b--noise measurement

G36.101 General Test Conditions.
G36.103 Acoustical measurement system.
G36.105 Sensing, recording, and reproducing equipment.
G36.107 Noise measurement procedures.
G36.109 Data recording, reporting, and approval.
G36.111 Flight procedures.

                        part c--data corrections

G36.201 Corrections to Test Results.
G36.203 Validity of results.

                          part d--noise limits

G36.301 Aircraft Noise Limits.

                             part a--general

    Section G36.1 Scope. This appendix prescribes limiting noise levels 
and procedures for measuring noise and adjusting these data to standard 
conditions, for propeller driven small airplanes and propeller-driven, 
commuter category airplanes specified in Sec. Sec. 36.1 and 36.501(c).

                        part b--noise measurement

                  Sec. G36.101 General Test Conditions.

    (a) The test area must be relatively flat terrain having no 
excessive sound absorption characteristics such as those caused by 
thick, matted, or tall grass, by shrubs, or by wooded areas. No 
obstructions which significantly influence the sound field from the 
airplane may exist within a conical space above the measurement 
position, the cone being defined by an axis normal to the ground and by 
a half-angle 75 degrees from the normal ground axis.
    (b) The tests must be carried out under the following conditions:
    (1) No precipitation;
    (2) Ambient air temperature between 36 and 95 degrees F (2.2 and 35 
degrees C);
    (3) Relative humidity between 20 percent and 95 percent, 
inclusively;
    (4) Wind speed may not exceed 10 knots (19 km/h) and cross wind may 
not exceed 5 knots (9 km/h), using a 30-second average;
    (5) No temperature inversion or anomalous wind condition that would 
significantly alter the noise level of the airplane when the nose is 
recorded at the required measuring point, and
    (6) The meteorological measurements must be made between 4 ft. (1.2 
m) and 33 ft. (10 m) above ground level. If the measurement site is 
within 1 n.m. of an airport meteorological station, measurements from 
that station may be used.
    (c) The flight test procedures, measuring equipment, and noise 
measurement procedures must be approved by the FAA.
    (d) Sound pressure level data for noise evaluation purposes must be 
obtained with acoustical equipment that complies with section G36.103 of 
this appendix.

               Sec. G36.103 Acoustical Measurement System.

    The acoustical measurement system must consist of approved equipment 
with the following characteristics: (a) A microphone system with 
frequency response compatible with measurement and analysis system 
accuracy as prescribed in section G36.105 of this appendix.
    (b) Tripods or similar microphone mountings that minimize 
interference with the sound being measured.
    (c) Recording and reproducing equipment characteristics, frequency 
response, and dynamic range compatible with the response and accuracy 
requirements of section G36.105 of this appendix.
    (d) Acoustic calibrators using sine wave or broadband noise of known 
sound pressure level. If broadband noise is used, the signal must be 
described in terms of its average and maximum root-mean-square (rms) 
value for non-overload signal level.

[[Page 984]]

       Sec. G36.105 Sensing, Recording, and Reproducing Equipment.

    (a) The noise produced by the airplane must be recorded. A magnetic 
tape recorder, graphic level recorder, or sound level meter is 
acceptable when approved by the regional certificating authority.
    (b) The characteristics of the complete system must comply with the 
requirements in International Electrotechnical Commission (IEC) 
Publications No. 651, entitled ``Sound Level Meters'' and No. 561, 
entitled ``Electro-acoustical Measuring Equipment for Aircraft Noise 
Certification'' as incorporated by reference under Sec. 36.6 of this 
part. Sound level meters must comply with the requirements for Type 1 
sound level meters as specified in IEC Publication No. 651.
    (c) The response of the complete system to a sensibly plane 
progressive sinusoidal wave of constant amplitude must be within the 
tolerance limits specified in IEC Publication No. 651, over the 
frequency range 45 to 11,200 Hz.
    (d) If equipment dynamic range limitations make it necessary, high 
frequency pre-emphasis must be added to the recording channel with the 
converse de-emphasis on playback. The pre-emphasis must be applied such 
that the instantaneous recorded sound pressure level of the noise signal 
between 800 and 11,200 Hz does not vary more than 20 dB between the 
maximum and minimum one-third octave bands.
    (e) The output noise signal must be read through an ``A'' filter 
with dynamic characteristics designated ``slow'' as defined in IEC 
Publication No. 651. A graphic level recorder, sound level meter, or 
digital equivalent may be used.
    (f) The equipment must be acoustically calibrated using facilities 
for acoustic free-field calibration and if analysis of the tape 
recording is requested by the Administrator, the analysis equipment 
shall be electronically calibrated by a method approved by the FAA. 
Calibrations shall be performed, as appropriate, in accordance with 
paragraphs A36.3.8 and A36.3.9 of appendix A of this part.
    (g) A windscreen must be employed with the microphone during all 
measurements of aircraft noise when the wind speed is in excess of 5 
knots (9 km/hr).

               Sec. G36.107 Noise Measurement Procedures.

    (a) The microphone must be a pressure type, 12.7 mm in diameter, 
with a protective grid, mounted in an inverted position such that the 
microphone diaphragm is 7 mm above and parallel to a white-painted metal 
circular plate. This white-painted metal plate shall be 40 cm in 
diameter and at least 2.5 mm thick. The plate shall be placed 
horizontally and flush with the surrounding ground surface with no 
cavities below the plate. The microphone must be located three-quarters 
of the distance from the center to the back edge of the plate along a 
radius normal to the line of flight of the test airplane.
    (b) Immediately prior to and after each test, a recorded acoustic 
calibration of the system must be made in the field with an acoustic 
calibrator for the purposes of checking system sensitivity and providing 
an acoustic reference level for the analysis of the sound level data. If 
a tape recorder or graphic level recorder is used, the frequency 
response of the electrical system must be determined at a level within 
10 dB of the full-scale reading used during the test, utilizing pink or 
pseudorandom noise.
    (c) The ambient noise, including both acoustic background and 
electrical systems noise, must be recorded and determined in the test 
area with the system gain set at levels which will be used for aircraft 
noise measurements. If aircraft sound pressure levels do not exceed the 
background sound pressure levels by at least 10 dB(A), a takeoff 
measurement point nearer to the start of the takeoff roll must be used 
and the results must be adjusted to the reference measurement point by 
an approved method.

          Sec. G36.109 Data Recording, Reporting, and Approval.

    (a) Data representing physical measurements and adjustments to 
measured data must be recorded in permanent form and appended to the 
record, except that corrections to measurements for normal equipment 
response deviations need not be reported. All other adjustments must be 
approved. Estimates must be made of the individual errors inherent in 
each of the operations employed in obtaining the final data.
    (b) Measured and corrected sound pressure levels obtained with 
equipment conforming to the specifications in section G36.105 of this 
appendix must be reported.
    (c) The type of equipment used for measurement and analysis of all 
acoustical, airplane performance, and meteorological data must be 
reported.
    (d) The following atmospheric data, measured immediately before, 
after, or during each test at the observation points prescribed in 
section G36.101 of this appendix must be reported:
    (1) Ambient temperature and relative humidity.
    (2) Maximum and average wind speeds and directions for each run.
    (e) Comments on local topography, ground cover, and events that 
might interfere with sound recordings must be reported.
    (f) The aircraft position relative to the takeoff reference flight 
path must be determined by an approved method independent of normal 
flight instrumentation, such as radar tracking, theodolite 
triangulation, or photographic scaling techniques.

[[Page 985]]

    (g) The following airplane information must be reported:
    (1) Type, model, and serial numbers (if any) of airplanes, engines, 
and propellers;
    (2) Any modifications or nonstandard equipment likely to affect the 
noise characteristics of the airplane;
    (3) Maximum certificated takeoff weight;
    (4) For each test flight, airspeed and ambient temperature at the 
flyover altitude over the measuring site determined by properly 
calibrated instruments;
    (5) For each test flight, engine performance parameters, such as 
manifold pressure or power, propeller speed (rpm) and other relevant 
parameters. Each parameter must be determined by properly calibrated 
instruments. For instance, propeller RPM must be validated by an 
independent device accurate to within 1 percent, 
when the airplane is equipped with a mechanical tachometer.
    (6) Airspeed, position, and performance data necessary to make the 
corrections required in section G36.201 of this appendix must be 
recorded by an approved method when the airplane is directly over the 
measuring site.

                     Sec. G36.111 Flight Procedures.

    (a) The noise measurement point is on the extended centerline of the 
runway at a distance of 8200 ft (2500 m) from the start of takeoff roll. 
The aircraft must pass over the measurement point within 10 degrees from the vertical and within 20% of the 
reference altitude. The flight test program shall be initiated at the 
maximum approved takeoff weight and the weight shall be adjusted back to 
this maximum weight after each hour of flight time. Each flight test 
must be conducted at the speed for the best rate of climb 
(Vy) 5 knots (9 
km/hour) indicated airspeed. All test, measurement, and data correction 
procedures must be approved by the FAA.
    (b) The takeoff reference flight path must be calculated for the 
following atmospheric conditions:
    (1) Sea level atmospheric pressure of 1013.25 mb (013.25 hPa);
    (2) Ambient air temperature of 59 [deg]F (15 [deg]C);
    (3) Relative humidity of 70 percent; and
    (4) Zero wind.
    (c) The takeoff reference flight path must be calculated assuming 
the following two segments:
    (1) First segment.
    (i) Takeoff power must be used from the brake release point to the 
point at which the height of 50 ft (15m) above the runway is reached.
    (ii) A constant takeoff configuration selected by the applicant must 
be maintained through this segment.
    (iii) The maximum weight of the airplane at brake-release must be 
the maximum for which noise certification is requested.
    (iv) The length of this first segment must correspond to the 
airworthiness approved value for a takeoff on a level paved runway (or 
the corresponding value for seaplanes).
    (2) Second segment.
    (i) The beginning of the second segment corresponds to the end of 
the first segment.
    (ii) The airplane must be in the climb configuration with landing 
gear up, if retractable, and flap setting corresponding to normal climb 
position throughout this second segment.
    (iii) The airplane speed must be the speed for the best rate of 
climb (Vy).
    (iv) For airplanes equipped with fixed pitch propellers, takeoff 
power must be maintained throughout the second segment. For airplanes 
equipped with variable pitch or constant speed propellers, takeoff power 
and rpm must be maintained throughout the second segment. If 
airworthiness limitations do not allow the application of takeoff power 
and rpm up to the reference point, then takeoff power and rpm must be 
maintained for as long as is permitted by such limitations; thereafter, 
maximum continuous power and rpm must be maintained. Maximum time 
allowed at takeoff power under the airworthiness standards must be used 
in the second segment. The reference height must be calculated assuming 
climb gradients appropriate to each power setting used.

                        part c--data corrections

                Sec. G36.201 Corrections to Test Results.

    (a) These corrections account for the effects of:
    (1) Differences in atmospheric absorption of sound between 
meteorological test conditions and reference conditions.
    (2) Differences in the noise path length between the actual airplane 
flight path and the reference flight path.
    (3) The change in the helical tip Mach number between test and 
reference conditions.
    (4) The change in the engine power between test and reference 
conditions.
    (b) Atmospheric absorption correction is required for noise data 
obtained when the test conditions are outside those specified in Figure 
G1. Noise data outside the applicable range must be corrected to 59 F 
and 70 percent relative humidity by an FAA approved method.

[[Page 986]]

[GRAPHIC] [TIFF OMITTED] TR13OC99.001

    (c) No corrections for helical tip Mach number variation need to be 
made if the propeller helical tip Mach number is:
    (1) At or below 0.70 and the test helical tip Mach number is within 
0.014 of the reference helical tip Mach number.
    (2) Above 0.70 and at or below 0.80 and the test helical tip Mach 
number is within 0.007 of the reference helical tip Mach number.
    (3) Above 0.80 and the test helical tip Mach number is within 0.005 
of the reference helical tip Mach number. For mechanical tachometers, if 
the helical tip Mach number is above 0.8 and the test helical tip Mach 
number is within 0.008 of the reference helical tip Mach number.
    (d) When the test conditions are outside those specified, 
corrections must be applied by an approved procedure or by the following 
simplified procedure:
    (1) Measured sound levels must be corrected from test day 
meteorological conditions to reference conditions by adding an increment 
equal to
Delta (M)=(HT [alpha]--0.7 HR)/1000

where HT is the height in feet under test conditions, 
          HR is the height in feet under reference conditions 
          when the aircraft is directly over the noise measurement point 
          and [alpha] is the rate of absorption for the test day 
          conditions at 500 Hz as specified in SAE ARP 866A, entitled 
          ``Standard Values of Atmospheric Absorption as a function of 
          Temperature and Humidity for use in Evaluating Aircraft 
          Flyover Noise'' as incorporated by reference under Sec. 36.6.
    (2) Measured sound levels in decibels must be corrected for height 
by algebraically adding an increment equal to Delta (1). When test day 
conditions are within those specified in figure G1:

    Delta (1)=22 log (HT/HR)

where HT is the height of the test aircraft when directly 
          over the noise measurement point and HR is the 
          reference height.
    When test day conditions are outside those specified in figure G1:
    Delta (1)=20 log (HT/HR)

    (3) Measured sound levels in decibels must be corrected for helical 
tip Mach number by algebraically adding an increment equal to:
    Delta (2)=k log (MR/MT)

where MT and MR are the test and reference helical 
          tip Mach numbers, respectively. The constant ``k'' is equal to 
          the slope of the line obtained for measured values of the 
          sound level in dB(A) versus helical tip Mach number. The value 
          of k may be determined from approved data. A nominal value of 
          k=150 may be used when MT is smaller than 
          MR. No correction may be

[[Page 987]]

          made using the nominal value of k when MT is larger 
          than MR. The reference helical tip Mach number 
          MR is the Mach number corresponding to the 
          reference conditions (RPM, airspeed, temperature) above the 
          measurement point.
    (4) Measured sound levels in decibels must be corrected for engine 
power by algebraically adding an increment equal to

Delta (3)=K3 log (PR/PT)

where PR and PT are the test and reference engine 
          powers respectively obtained from the manifold pressure/torque 
          gauges and engine rpm. The value of K3 shall be 
          determined from approved data from the test airplane. In the 
          absence of flight test data and at the discretion of the 
          Administrator, a value of K3=17 may be used.

                    Sec. G36.203 Validity of Results.

    (a) The measuring point must be overflown at least six times. The 
test results must produce an average noise level (LAmax) 
value within a 90 percent confidence limit. The average noise level is 
the arithmetic average of the corrected acoustical measurements for all 
valid test runs over the measuring point.
    (b) The samples must be large enough to establish statistically a 90 
percent confidence limit not exceeding 1.5 dB(A). 
No test results may be omitted from the averaging process unless 
omission is approved by the FAA.

                          part d--noise limits

                   Sec. G36.301 Aircraft noise limits.

    (a) Compliance with this section must be shown with noise data 
measured and corrected as prescribed in Parts B and C of this appendix.
    (b) For single-engine airplanes for which the original type 
certification application is received before February 3, 2006 and multi-
engine airplanes, the noise level must not exceed 76 dB(A) up to and 
including aircraft weights of 1,320 pounds (600 kg). For aircraft 
weights greater than 1,320 pounds, the limit increases from that point 
with the logarithm of airplane weight at the rate of 9.83 dB (A) per 
doubling of weight, until the limit of 88 dB (A) is reached, after which 
the limit is constant up to and including 19,000 pounds (8,618 kg). 
Figure G2 shows noise level limits vs airplane weight.
    (c) For single-engine airplanes for which the original type 
certification application is received on or after February 3, 2006, the 
noise level must not exceed 70dB(A) for aircraft having a maximum 
certificated takeoff weight of 1,257 pounds (570 kg) or less. For 
aircraft weights greater than 1,257 pounds, the noise limit increases 
from that point with the logarithm of airplane weight at the rate of 
10.75dB(A) per doubling of weight, until the limit of 85dB(A) is 
reached, after which the limit is constant up to and including 19,000 
pounds (8,618 kg). Figure G2 depicts noise level limits for airplane 
weights for single-engine airplanes.
[GRAPHIC] [TIFF OMITTED] TR04JA06.033


[[Page 988]]



(Secs. 313(a), 603, and 611(b), Federal Aviation Act of 1958 as amended 
(49 U.S.C. 1354(a), 1423, and 1431(b)); sec. 6(c), Department of 
Transportation Act (49 U.S.C. 1655 (c)); Title I, National Environmental 
Policy Act of 1969 (42 U.S.C. 4321 et seq.); E. O. 11514, March 5, 1970 
and 14 CFR 11.45).

[Amdt. 36-16, 53 FR 47400, Nov. 22, 1988; 53 FR 50157, Dec. 13, 1988, as 
amended by Amdt. 36-22, 64 FR 55602, Oct. 13, 1999; Amdt. 36-54, 67 FR 
45236, July 8, 2002; Amdt. 36-27, 70 FR 45504, Aug. 5, 2005; Amdt. 36-
28, 71 FR 532, Jan. 4, 2006]



  Sec. Appendix H to Part 36--Noise Requirements For Helicopters Under 
                                Subpart H

                      part a--reference conditions

Sec.

H36.1 General.
H36.3 Reference Test Conditions.
H36.5 Symbols and Units.

              part b--noise measurement under Sec. 36.801

H36.101 Noise certification test and measurement conditions.
H36.103 Takeoff test conditions.
H36.105 Flyover test conditions.
H36.107 Approach test conditions.
H36.109 Measurement of helicopter noise received on the ground.
H36.111 Reporting and correcting measured data.
H36.113 Atmospheric attenuation of sound.

       part c--noise evaluation and calculation under Sec. 36.803

H36.201 Noise evaluation in EPNdB.
H36.203 Calculation of noise levels.
H36.205 Detailed data correction procedures.

                 part d--noise limits under Sec. 36.805

H36.301 Noise measurement, evaluation, and calculation.
H36.303 [Reserved]
H36.305 Noise levels.

                      part a--reference conditions

    Section H36.1 General. This appendix prescribes noise requirements 
for helicopters specified under Sec. 36.1, including:
    (a) The conditions under which helicopter noise certification tests 
under Part H must be conducted and the measurement procedures that must 
be used under Sec. 36.801 to measure helicopter noise during each test;
    (b) The procedures which must be used under Sec. 36.803 to correct 
the measured data to the reference conditions and to calculate the noise 
evaluation quantity designated as Effective Perceived Noise Level 
(EPNL); and
    (c) The noise limits for which compliance must be shown under Sec. 
36.805.

                Section H36.3 Reference Test Conditions.

    (a) Meteorological conditions. Aircraft position, performance data 
and noise measurements must be corrected to the following noise 
certification reference atmospheric conditions which shall be assumed to 
exist from the surface to the aircraft altitude:
    (1) Sea level pressure of 2,116 psf (1,013.25 hPa).
    (2) Ambient temperature of 77 degrees F (25 degrees C).
    (3) Relative humidity of 70 percent.
    (4) Zero wind.
    (b) Reference test site. The reference test site is flat and without 
line-of-sight obstructions across the flight path that encompasses the 
10 dB down points.
    (c) Takeoff reference profile. (1) Figure H1 illustrates a typical 
takeoff profile, including reference conditions.
    (2) The reference flight path is defined as a straight line segment 
inclined from the starting point (1,640 feet (500 meters) from the 
center microphone location and 65 feet (20 meters) above ground level) 
at a constant climb angle [beta] defined by the certificated best rate 
of climb and Vy for minimum engine performance. The constant 
climb angle [beta] is derived from the manufacturer's data (approved by 
the FAA) to define the flight profile for the reference conditions. The 
constant climb angle [beta] is drawn through Cr and 
continues, crossing over station A, to the position corresponding to the 
end of the type certification takeoff path represented by position 
Ir.
    (d) Level flyover reference profile. The beginning of the level 
flyover reference profile is represented by helicopter position 
Dr (Figure H2). The helicopter approaches position 
Dr in level flight 492 feet above ground level as measured at 
Station A. Reference airspeed must be either 0.9VH; 
0.9VNE; 0.45VH + 65 kts (0.45VH + 
120km/h); or 0.45VNE + 65kts (0.45VNE + 120 km/h), 
whichever of the four speeds is least. The helicopter crosses directly 
overhead station A in level flight and proceeds to position 
Jr.
    (e) For noise certification purposes, VH is defined as 
the airspeed in level flight obtained using the minimum specified engine 
torque corresponding to maximum continuous power available for sea level 
pressure of 2,116 psf (1,013.25 hPa) at 77 [deg]F (25 [deg]C) ambient 
conditions at the relevant maximum certificated weight. The value of 
VNE is the never-exceed airspeed. The values of VH 
and VNE that are used for noise certification must be listed 
in the approved Rotorcraft Flight Manual.
    (f) Approach reference profile. (1) Figure H3 illustrates approach 
profile, including reference conditions.

[[Page 989]]

    (i) The beginning of the approach profile is represented by 
helicopter position E. The position of the helicopter is recorded for a 
sufficient distance (EK) to ensure recording of the entire interval 
during which the measured helicopter noise level is within 10 dB of 
Maximum Tone Corrected Perceived Noise Level (PNLTM). The reference 
flight path, ErKr represents a stable flight 
condition in terms of torque, rpm, indicated airspeed, and rate of 
descent resulting in a 6[deg] approach angle.
    (ii) The test approach profile is defined by the approach angle 
[eta] passing directly over the station A at a height of AH, to position 
K, which terminates the approach noise certification profile. The test 
approach angle [eta] must be between 5.5[deg] and 6.5[deg].
    (2) The helicopter approaches position H along a constant 6[deg] 
approach slope throughout the 10 dB down time period. The helicopter 
crosses position E and proceeds along the approach slope crossing over 
station A until it reaches position K.

    Section H36.5 Symbols and units. The following symbols and units as 
used in this appendix for helicopter noise certification have the 
following meanings.

                Flight Profile Identification--Positions
------------------------------------------------------------------------
         Position                            Description
------------------------------------------------------------------------
A.........................  Location of the noise measuring point at the
                             flight-track noise measuring station
                             vertically below the reference (takeoff,
                             flyover, or approach) flight path.
C.........................  Start of noise certification takeoff flight
                             path.
Cr........................  Start of noise certification reference
                             takeoff flight path.
D.........................  Start of noise certification flyover flight
                             path.
Dr........................  Start of noise certification reference
                             flyover path.
E.........................  Start of noise certification approach flight
                             path.
Er........................  Start of noise certification reference
                             approach flight path.
F.........................  Position on takeoff flight path directly
                             above noise measuring station A.
Fr........................  Position on reference takeoff path directly
                             above noise measuring Station A.
G.........................  Position on flyover flight path directly
                             above noise measuring station A.
Gr........................  Position on reference flyover path directly
                             above noise measuring Station A.
H.........................  Position on approach flight path directly
                             above noise measuring station A.
Hr........................  Position on reference path directly above
                             noise measuring Station A.
I.........................  End of noise type certification takeoff
                             flight path.
Ir........................  End of noise type certification reference
                             takeoff flight path.
J.........................  End of noise type certification flyover
                             flight path.
Jr........................  End of noise type certification reference
                             flyover flight path.
K.........................  End of noise certification approach type
                             flight path.
Kr........................  End of noise type certification reference
                             approach flight path.
L.........................  Position on measured takeoff flight path
                             corresponding to PNLTM at station A.
Lr........................  Position on reference takeoff flight path
                             corresponding to PNLTM of station A.
M.........................  Position on measured flyover flight path
                             corresponding to PNLTM of station A.
Mr........................  Position on reference flyover flight path
                             corresponding to PNLTM of station A.
N.........................  Position on measured approach flight path
                             corresponding to PNLTM at station A.
Nr........................  Position on reference approach flight path
                             corresponding to PNLTM at station A.
S.........................  Sideline noise measuring station (note: a
                             subscript denotes the aircraft orientation
                             relative to the direction of flight).
------------------------------------------------------------------------


                        Flight Profile Distances
------------------------------------------------------------------------
      Distance             Unit                     Meaning
------------------------------------------------------------------------
AF.................  Feet...........  Takeoff Height. The vertical
                                       distance between helicopter and
                                       station A.
AG.................  Feet...........  Flyover Height. The vertical
                                       distance between the helicopter
                                       and station A.
AH.................  Feet...........  Approach Height. The vertical
                                       distance between the helicopter
                                       and station A.
AL.................  Feet...........  Measured Takeoff Noise Path. The
                                       distance from station A to the
                                       measured helicopter position L.
ALr................  Feet...........  Reference Takeoff Noise Path. The
                                       distance from station A to the
                                       reference helicopter position Lr.
AM.................  Feet...........  Measured Flyover Noise Path. The
                                       distance from station A to the
                                       measured helicopter position M.
AMr................  Feet...........  Reference Flyover Noise Path. The
                                       distance from station A to
                                       helicopter position Mr on the
                                       reference flyover flight path.
AN.................  Feet...........  Measured Approach Noise Path. The
                                       distance from station A to the
                                       measured helicopter noise
                                       position N.
ANr................  Feet...........  Reference Approach Noise Path. The
                                       distance from station A to the
                                       reference helicopter position Nr.
CI.................  Feet...........  Takeoff Flight Path Distance. The
                                       distance from position C at which
                                       the helicopter establishes a
                                       constant climb angle on the
                                       takeoff flight path passing over
                                       station A and continuing to
                                       position I at which the position
                                       of the helicopter need no longer
                                       be recorded.
DJ.................  Feet...........  Flyover Flight Path Distance. The
                                       distance from position D at which
                                       the helicopter is established on
                                       the flyover flight path passing
                                       over station A and continuing to
                                       position J at which the position
                                       of the helicopter need no longer
                                       be recorded.
EK.................  Feet...........  Approach Flight Path Distance. The
                                       distance from position E at which
                                       the helicopter establishes a
                                       constant angle on the approach
                                       flight path passing over station
                                       A and continuing to position K at
                                       which the position of the
                                       helicopter need no longer be
                                       recorded.
------------------------------------------------------------------------

              part b--noise measurement under Sec. 36.801

  Section H36.101 Noise certification test and measurement conditions.

    (a) General. This section prescribes the conditions under which 
aircraft noise certification tests must be conducted and the

[[Page 990]]

measurement procedures that must be used to measure helicopter noise 
during each test.
    (b) Test site requirements. (1) Tests to show compliance with 
established helicopter noise certification levels must consist of a 
series of takeoffs, level flyovers, and approaches during which 
measurement must be taken at noise measuring stations located at the 
measuring points prescribed in this section.
    (2) Each takeoff test, flyover test, and approach test includes 
simultaneous measurements at the flight-track noise measuring station 
vertically below the reference flight path and at two sideline noise 
measuring stations, one on each side of the reference flight track 492 
feet (150m) from, and on a line perpendicular to, the flight track of 
the noise measuring station.
    (3) The difference between the elevation of either sideline noise 
measuring station may not differ from the flight-track noise measuring 
station by more than 20 feet.
    (4) Each noise measuring station must be surrounded by terrain 
having no excessive sound absorption characteristics, such as might be 
caused by thick, matted, or tall grass, shrubs, or wooded areas.
    (5) During the period when the takeoff, flyover, or approach noise/
time record indicates the noise measurement is within 10 dB of PNLTM, no 
obstruction that significantly influences the sound field from the 
aircraft may exist--
    (i) For any flight-track or sideline noise measuring station, within 
a conical space above the measuring position (the point on the ground 
vertically below the microphone), the cone being defined by an axis 
normal to the ground and by half-angle 80[deg] from this axis; and
    (ii) For any sideline noise measuring station, above the line of 
sight between the microphone and the helicopter.
    (6) If a takeoff or flyover test series is conducted at weights 
other than the maximum takeoff weight for which noise certification is 
requested, the following additional requirements apply:
    (i) At least one takeoff test and one flyover test must be conducted 
at, or above, the maximum certification weight.
    (ii) Each test weight must be within +5 percent or -10 percent of 
the maximum certification weight.
    (7) Each approach test must be conducted with the aircraft 
stabilized and following a 6.0 degree 0.5 degree 
approach angle and must meet the requirements of section H36.107 of this 
part.
    (8) If an approach test series is conducted at weights other than 
the maximum landing weight for which certification is requested, the 
following additional requirements apply:
    (i) At least one approach test must be conducted at a weight at, or 
above, the maximum landing weight.
    (ii) Each test weight must be between +5 percent and -10 percent of 
the maximum certification weight.
    (c) Weather restrictions. The tests must be conducted under the 
following atmospheric conditions:
    (1) No rain or other precipitation.
    (2) Ambient air temperature between 14 [deg]F and 95 [deg]F (-10 
[deg]C and 35 [deg]C), inclusively, at a point 33 feet (10 meters) above 
the ground at the noise measuring station and at the aircraft. The 
temperature and relative humidity measured at a point 33 feet (10 
meters) above the ground at the noise measuring station must be used to 
adjust for propagation path absorption.
    (3) Relative humidity and ambient temperature at a point 33 feet (10 
meters) above the ground at the noise measuring station and at the 
aircraft, is such that the sound attenuation in the one-third octave 
band centered at 8 kHz is not greater than 12 dB/100 meters and the 
relative humidity is between 20 percent and 95 percent, inclusively.
    (4) Wind velocity as measured at 10 meters above ground does not 
exceed 10 knots (19 km/h) and the crosswind component does not exceed 5 
knots (9 km/h). The wind shall be determined using a continuous thirty-
second averaging period spanning the 10dB down time interval.
    (5) No anomalous meteorological conditions (including turbulence) 
that will significantly affect the noise level of the aircraft when the 
noise is recorded at each noise measuring station.
    (6) The wind velocity, temperature, and relative humidity 
measurements required under the appendix must be measured in the 
vicinity of noise measuring stations 10 meters above the ground. The 
location of the meteorological measurements must be approved by the FAA 
as representative of those atmospheric conditions existing near the 
surface over the geographical area which aircraft noise measurements are 
made. In some cases, a fixed meteorological station (such as those found 
at airports or other facilities) may meet this requirement.
    (7) Temperature and relative humidity measurements must be obtained 
within 30 minutes of each noise test.
    (d) Aircraft testing procedures. (1) The aircraft testing procedures 
and noise measurements must be conducted and processed in a manner that 
yields the noise evaluation measure designated as Effective Perceived 
Noise Level (EPNL) in units of EPNdB, as prescribed in Appendix A of 
this part.
    (2) The helicopter height and lateral position relative to the 
reference flight track (which passes through the flight track noise 
measuring station) must be determined using an FAA-approved method. The 
equipment used to make the determination must be independent of normal 
flight instrumentation. Applicable independent systems are

[[Page 991]]

radar tracking, theodolite triangulation, laser trajectography, photo 
scaling, or differential global positioning system.
    (3) The helicopter position along the flight path must be related to 
the noise recorded at the noise measuring stations by means of 
synchronized signals recorded at an approved sampling rate. The 
helicopter position must be recorded relative to the reference flight 
track during the entire time interval in which the recorded signal is 
within 10 dB of PNLTM. Measuring and sampling equipment must be approved 
by the FAA before testing.
    (4) Aircraft performance data sufficient to make the corrections 
required under section H36.205 of this appendix must be recorded at an 
FAA-approved sampling rate using FAA-approved equipment.

                Section H36.103 Takeoff test conditions.

    (a) This section, in addition to the applicable requirements of 
sections H36.101 and H36.205(b) of this appendix, applies to all takeoff 
noise tests conducted under this appendix to show compliance with Part 
36.
    (b) A test series must consist of at least six flights over the 
flight-track noise measuring station (with simultaneous measurements at 
all three noise measuring stations) as follows:
    (1) An airspeed of either Vy 5 
knots or the lowest approved speed 5 knots for the 
climb after takeoff, whichever speed is greater, must be established and 
maintained throughout the 10 dB-down time interval.
    (2) The horizontal portion of each test flight must be conducted at 
an altitude of 65 feet (20 meters) above the ground level at the flight-
track noise measuring station.
    (3) Upon reaching a point 1,640 feet (500 meters) from the noise 
measuring station, the helicopter must be stabilized at the maximum 
takeoff power that corresponds to minimum installed engine(s) 
specification power available for the reference ambient conditions or 
gearbox torque limit, whichever is lower.
    (4) The helicopter must be maintained throughout the 10 dB-down time 
interval at the best rate of climb speed Vy 5 knots, or the lowest approved speed for climb after 
takeoff, whichever is greater, for an ambient temperature of 25 [deg]C 
at sea level.
    (5) The average rotor speed must not vary from the maximum normal 
operating rotor RPM by more than 1.0 percent 
during the 10 dB-down time interval.
    (6) The helicopter must stay within 10[deg] or 
65 feet (20 meters), 
whichever is greater, from the vertical above the reference track 
throughout the 10dB-down time interval.
    (7) A constant takeoff configuration selected by the applicant must 
be maintained throughout the takeoff reference procedure with the 
landing gear position consistent with the airworthiness certification 
tests for establishing best rate-of-climb speed, Vy.

                Section H36.105 Flyover test conditions.

    (a) This section, in addition to the applicable requirements of 
sections H36.101 and H36.205(c) of this appendix, applies to all flyover 
noise tests conducted under this appendix to show compliance with Part 
36.
    (b) A test series consists of at least six flights. The number of 
level flights made with a headwind component must be equal to the number 
of level flights made with a tailwind component with simultaneous 
measurements at all three noise measuring stations--
    (1) In level flight cruise configuration;
    (2) At a height of 492 feet 30 feet (150 
9 meters) above the ground level at the flight-
track noise measuring station; and
    (3) The helicopter must fly within 10[deg] or 
65 feet (20 meters), 
whichever is greater, from the vertical above the reference track 
throughout the 10 dB-down time interval.
    (c) Each flyover noise test must be conducted--
    (1) At a speed of 0.9VH; 0.9VNE; 
0.45VH + 65 kts (0.45VH + 120 km/h); or 
0.45VNE + 65 kts (0.45VNE + 120 km/h), whichever 
speed is least, to be maintained throughout the measured portion of the 
flyover;
    (2) At average rotor speed, which must not vary from the maximum 
normal operating rotor RPM by more than 1.0 
percent during the 10 dB-down time interval.
    (3) With the power stabilized during the period when the measured 
helicopter noise level is within 10 dB of PNLTM.
    (d) The airspeed shall not vary from the reference airspeed by more 
than 5 knots (9 km/hr).

                Section H36.107 Approach test conditions.

    (a) This section, in addition to the requirements of sections 
H36.101 and H36.205(d) of this appendix, applies to all approach tests 
conducted under this appendix to show compliance with Part 36.
    (b) A test series must consist of at least six flights over the 
flight-track noise measuring station (with simultaneous measurements at 
the three noise measuring stations)--
    (1) On an approach slope of 6[deg] 0.5[deg];
    (2) At a height of 394 33 feet (120 10 meters)
    (3) The helicopter must fly within 10[deg] or 
65 feet (20 meters) lateral 
deviation tolerance, whichever is greater, from the vertical above the 
reference track throughout the 10 dB-down time interval;
    (4) At stabilized airspeed equal to the certificated best rate of 
climb Vy, or the lowest approved speed for approach, 
whichever is greater, with power stabilized during the approach and over 
the flight path reference point, and continued to a normal touchdown; 
and
    (5) At average rotor speed, which may not vary from the maximum 
normal operating

[[Page 992]]

rotor RPM by more than 1.0 percent during the 10 
dB-down time interval; and
    (6) The constant approach configuration used in airworthiness 
certification tests, with the landing gear extended, must be maintained 
throughout the approach reference procedure.
    (c) The airspeed shall not vary from the reference airspeed by more 
than 5 knots (9 km/hr).

 Section H36.109 Measurement of Helicopter Noise Received on the Ground.

    The measurement system and the measurement, calibration and general 
analysis procedures to be used are provided in Appendix A, section A36.3 
of this part.

         Section H36.111 Reporting and correcting measured data.

    (a) General. Data representing physical measurements, and 
corrections to measured data, including corrections to measurements for 
equipment response deviations, must be recorded in permanent form and 
appended to the record. Each correction must be reported and is subject 
to FAA approval. An estimate must be made of each individual error 
inherent in each of the operations employed in obtaining the final data.
    (b) Data reporting. (1) Measured and corrected sound pressure levels 
must be presented in one-third octave band levels obtained with 
equipment conforming to the standards prescribed in section H36.109 of 
this appendix.
    (2) The type of equipment used for measurement and analysis of all 
acoustic, aircraft performance, and meteorological data must be 
reported.
    (3) The atmospheric environmental data required to demonstrate 
compliance with this appendix, measured throughout the test period, must 
be reported.
    (4) Conditions of local topography, ground cover, or events which 
may interfere with sound recording must be reported.
    (5) The following aircraft information must be reported:
    (i) Type, model, and serial numbers, if any, of aircraft engines and 
rotors.
    (ii) Gross dimensions of aircraft and location of engines.
    (iii) Aircraft gross weight for each test run.
    (iv) Aircraft configuration, including landing gear positions.
    (v) Airspeed in knots.
    (vi) Helicopter engine performance as determined from aircraft 
instruments and manufacturer's data.
    (vii) Aircraft flight path, above ground level in feet, determined 
by an FAA approved method which is independent of normal flight 
instrumentation, such as radar tracking, theodolite triangulation, laser 
trajectography, or photographic scaling techniques.
    (6) Aircraft speed, and position, and engine performance parameters 
must be recorded at an approved sampling rate sufficient to correct to 
the noise certification reference test conditions prescribed in section 
H36.3 of this appendix. Lateral position relative to the reference 
flight-track must be reported.
    (c) Data corrections. (1) Aircraft position, performance data and 
noise measurement must be corrected to the noise certification reference 
conditions as prescribed in sections H36.3 and H36.205 of this appendix.
    (2) The measured flight path must be corrected by an amount equal to 
the difference between the applicant's predicted flight path for the 
certification reference conditions and the measured flight path at the 
test conditions. Necessary corrections relating to helicopter flight 
path or performance may be derived from FAA-approved data for the 
difference between measured and reference conditions, together with 
appropriate allowances for sound attenuation with distance. The 
Effective Perceived Noise Level (EPNL) correction may not exceed 2.0 
EPNdB except for takeoff flight condition, where the correction may not 
exceed 4.0 EPNdB, of which the arithmetic sum of [Delta]1 
(described in section H36.205(f)(1)) and the term -7.5 log (AL/
ALr) from [Delta]2 term (described in section 
H36.205(g)(1)(i)) may not exceed 2.0 EPNdB, for any combination of the 
following:
    (i) The helicopter not passing vertically above the measuring 
station.
    (ii) Any difference between the reference flight track and the 
actual test flight track; and
    (iii) Detailed correction requirements prescribed in section H36.205 
of this appendix.
    (3) Helicopter sound pressure levels within the 10 dB-down time 
interval must exceed the mean background sound pressure levels 
determined under section B36.3.9.11 by at least 3 dB in each one-third 
octave band, or must be corrected under an FAA-approved method.
    (d) Validity of results. (1) The test results must produce three 
average EPNL values within the 90 percent confidence limits, each value 
consisting of the arithmetic average of the corrected noise measurements 
for all valid test runs at the takeoff, level flyovers, and approach 
conditions. The 90 percent confidence limit applies separately to 
takeoff, flyover, and approach.
    (2) The minimum sample size acceptable for each takeoff, approach, 
and flyover certification measurements is six. The number of samples 
must be large enough to establish statistically for each of the three 
average noise certification levels a 90 percent confidence limit which 
does not exceed 1.5 EPNdB. No test result may be 
omitted from the averaging process, unless otherwise specified by the 
FAA.

[[Page 993]]

    (3) To comply with this appendix, a minimum of six takeoffs, six 
approaches, and six level flyovers is required. To be counted toward 
this requirement, each flight event must be validly recorded at all 
three noise measuring stations.
    (4) The approved values of VH and Vy used in 
calculating test and reference conditions and flight profiles must be 
reported along with measured and corrected sound pressure levels.

            Section H36.113 Atmospheric attenuation of sound.

    (a) The values of the one-third octave band spectra measured during 
helicopter noise certification tests under this appendix must conform, 
or be corrected, to the reference conditions prescribed in section 
H36.3(a). Each correction must account for any differences in the 
atmospheric attenuation of sound between the test-day conditions and the 
reference-day conditions along the sound propagation path between the 
aircraft and the microphone. Unless the meteorological conditions are 
within the test window prescribed in this appendix, the test data are 
not acceptable.
    (b) Attenuation rates. The procedure for determining the atmospheric 
attenuation rates of sound with distance for each one-third octave bands 
must be determined in accordance with Society of Automotive Engineering 
(SAE) ARP 866A. The atmospheric attenuation equations are provided in 
both the International and English system of units in section A36.7 of 
this part.
    (c) Correction for atmospheric attenuation. (1) EPNL values 
calculated for measured data must be corrected whenever--
    (i) The ambient atmospheric conditions of temperature and relative 
humidity do not conform to the reference conditions, 77 [deg]F and 70%, 
respectively, or
    (ii) The measured flight paths do not conform to the reference 
flight paths.
    (iii) The temperature and relative humidity measured at 33 feet (10 
meters) above the ground must be used to adjust for propagation path 
absorption.
    (2) The mean attenuation rate over the complete sound propagation 
path from the aircraft to the microphone must be computed for each one-
third octave band from 50 Hz to 10,000 Hz. These rates must be used in 
computing the corrections required in section H36.111(d) of this 
appendix.

       part c--noise evaluation and calculation under Sec. 36.803

               Section H36.201 Noise Evaluation in EPNdB.

    (a) Effective Perceived Noise Level (EPNL), in units of effective 
perceived noise decibels (EPNdB), shall be used for evaluating noise 
level values under Sec. 36.803 of this part. Except as provided in 
paragraph (b) of this section, the procedures in appendix A of Part 36 
must be used for computing EPNL. appendix A includes requirements 
governing determination of noise values, including calculations of:
    (1) Perceived noise levels;
    (2) Corrections for spectral irregularities;
    (3) Tone corrections;
    (4) Duration corrections;
    (5) Effective perceived noise levels; and
    (6) Mathematical formulation of noy tables.
    (b) Notwithstanding the provisions of section A36.4.3.1(a), for 
helicopter noise certification, corrections for spectral irregularities 
shall start with the corrected sound pressure level in the 50 Hz one-
third octave band.

              Section H36.203 Calculation of noise levels.

    (a) To demonstrate compliance with the noise level limits of section 
H36.305, the noise values measured simultaneously at the three noise 
measuring points must be arithmetically averaged to obtain a single 
EPNdB value for each flight.
    (b) The calculated noise level for each noise test series, i.e., 
takeoff, flyover, or approach must be the numerical average of at least 
six separate flight EPNdB values. The 90 percent confidence limit for 
all valid test runs under section H36.111(d) of this appendix applies 
separately to the EPNdB values for each noise test series.

          Section H36.205 Detailed data correction procedures.

    (a) General. If the test conditions do not conform to those 
prescribed as noise certification reference conditions under section 
H36.305 of this appendix, the following correction procedure shall 
apply:
    (1) If there is any difference between measured test and reference 
conditions, an appropriate correction must be made to the EPNL 
calculated from the measured noise data. Conditions that can result in a 
different value include:
    (i) Atmospheric absorption of sound under measured test conditions 
that are different from the reference test conditions; or
    (ii) Measured flight path that is different from the reference 
flight path.
    (2) The following correction procedures may produce one or more 
possible correction values which must be added algebraically to the 
calculated EPNL to bring it to reference conditions:
    (i) The flight profiles must be determined for both reference and 
test conditions. The procedures require noise and flight path recording 
with a synchronized time signal from which the test profile can be 
delineated, including the aircraft position for which PNLTM is observed 
at the noise measuring station. For takeoff, the flight profile

[[Page 994]]

corrected to reference conditions may be derived from FAA approved 
manufacturer's data.
    (ii) The sound propagation paths to the microphone from the aircraft 
position corresponding to PNLTM must be determined for both the test and 
reference profiles. The SPL values in the spectrum of PNLTM must then be 
corrected for the effects of--
    (A) Change in atmospheric sound absorption;
    (B) Atmospheric sound absorption on the linear difference between 
the two sound path lengths; and
    (C) Inverse square law on the difference in sound propagation path 
length. The corrected values of SPL must then be converted to a 
reference condition PNLTM value from which PNLTM must be subtracted. The 
resulting difference represents the correction which must be added 
algebraically to the EPNL calculated from the measured data.
    (iii) As observed at the noise measuring station, the measured PNLTM 
distance is different from the reference PNLTM distance and therefore 
the ratio must be calculated and used to determine a noise duration 
correction factor. Effective perceived noise level, EPNL, is determined 
by the algebraic sum of the maximum tone corrected perceived noise level 
(PNLTM) and the duration correction factor.
    (iv) For aircraft flyover, alternative source noise corrections 
require FAA approval and must be determined and adjusted to account for 
noise level changes caused by the differences between measured test 
conditions and reference conditions.
    (b) Takeoff profiles. (1) Figure H1 illustrates a typical takeoff 
profile, including reference conditions.
    (i) The reference takeoff flight path is described in section 
H36.3(c).
    (ii) The test parameters are functions of the helicopter's 
performance and weight and the atmospheric conditions of temperature, 
pressure, wind velocity and direction.
    (2) For the actual takeoff, the helicopter approaches position C in 
level flight at 65 feet (20 meters) above ground level at the flight 
track noise measuring station and at either Vy 5 knots or the lowest approved speed for the climb after 
takeoff, whichever speed is greater.

[[Page 995]]

[GRAPHIC] [TIFF OMITTED] TR02JN04.000

    (3) Figure H1 illustrates the significant geometrical relationships 
influencing sound propagation. Position L represents the helicopter 
location on the measured takeoff path from which PNLTM is observed at 
station A, and Lr is the corresponding position on the 
reference sound propagation path. Propagation paths AL and 
ALr both form the same

[[Page 996]]

angle [thetas] (theta) relative to their respective flight paths.
    (c) Level flyover profiles. (1) The noise type certification level 
flyover profile is shown in Figure H2. Airspeed must be stabilized 
within 5 knots of the reference airspeed 
determined using the procedures in section H36.3(d). The number of level 
flights made with a headwind component must be equal to the number of 
level flights made with a tailwind component.

[[Page 997]]

[GRAPHIC] [TIFF OMITTED] TR02JN04.001

    (2) Figure H2 illustrates comparative flyover profiles when test 
conditions do not conform to prescribed reference conditions. The 
position of the helicopter shall be recorded for a distance (DJ) 
sufficient to ensure recording of the entire interval during which the 
measured helicopter noise level is

[[Page 998]]

within 10 dB of PNLTM, as required. The flyover profile is defined by 
the height AG which is a function of the operating conditions controlled 
by the pilot. Position M represents the helicopter location on the 
measured flyover flight path for which PNLTM is observed at station A, 
and Mr is the corresponding position on the reference flight 
path.
    (d) Approach profiles. (1) Figure H3 illustrates a typical approach 
profile, including reference conditions.
    (2) The helicopter approaches position H along a 6[deg] (0.5[deg]) average approach slope throughout the 10dB-
down time interval. Deviation from the 6[deg] average approach slope 
must be approved by the FAA before testing.

[[Page 999]]

[GRAPHIC] [TIFF OMITTED] TR02JN04.002

    (3) Figure H3 illustrates portions of the measured and reference 
approach flight paths including the significant geometrical 
relationships influencing sound propagation. The measured approach path 
is represented by segment EK with an approach allowable angle [thetas]. 
Reference positions, Er and Kr, define an 
idealized reference approach angle of 6[deg].

[[Page 1000]]

Position N represents the helicopter location on the measured approach 
flight path for which PNLTM is observed at measuring station A, and 
Nr is the corresponding position on the reference approach 
flight path. The measured and reference noise propagation paths are AN 
and ANr, respectively, both of which form the same angle, 
[thetas]APP, corresponding to PNLTM relative to their 
approach flight paths.
    (e) Correction of noise at source during level flyover. (1) For 
level overflight, if any combination of the following three factors, 
airspeed deviations from reference, rotor speed deviations from 
reference, and temperature deviations from reference, results in a noise 
correlating parameter whose value deviates from the reference value of 
this parameter, then source noise adjustments must be determined from 
the manufacturer's data that is approved by the FAA.
    (2) Off-reference tip Mach number adjustments must be based upon a 
sensitivity curve of PNLTM versus advancing blade tip Mach number, 
deduced from overflights performed at different airspeeds surrounding 
the reference airspeed. If the test aircraft is unable to attain the 
reference value, then an extrapolation of the sensitivity curve is 
permitted if data cover at least a range of 0.03 Mach units. The 
advancing blade tip Mach number must be computed using true airspeed, 
onboard outside air temperature, and rotor speed. A separate PNLTM 
versus advancing blade tip Mach number function must be derived for each 
of the three certification microphone locations, i.e., centerline, 
sideline left, and sideline right. Sideline left and right are defined 
relative to the direction of flight for each run. PNLTM adjustments are 
to be applied to each microphone datum using the appropriate PNLTM 
function.
    (f) PNLT corrections. If the measured ambient atmospheric conditions 
of temperature and relative humidity differ from those prescribed as 
reference conditions under this appendix (77 degrees F and 70 percent, 
respectively), corrections to the EPNL values must be calculated from 
the measured data under paragraph (a) of this section as follows:
    (1) Takeoff flight path. For the takeoff flight path shown in Figure 
H1, the spectrum of PNLTM observed at station A for the aircraft at 
position L is decomposed into its individual SPL(i) values.
    (i) Step 1. A set of corrected values are then computed as follows:

SPL(i)r = SPL(i) + C[[alpha](i) - 
          [alpha](i)o]AL + 
          C[alpha](i)o (AL - ALr) + 20 
          log (AL/ALr)

where SPL(i) and SPL(i)r are the measured and corrected sound 
pressure levels, respectively, in the i-th one-third octave band. The 
first correction term adjusts for the effect of change in atmospheric 
sound absorption where [alpha](i) and [alpha](i)o are the 
sound attenuation coefficients for the test and reference atmospheric 
conditions, respectively, for the i-th one-third octave band, and AL is 
the measured takeoff sound propagation path. The conversion factor 
constant, C, is 0.001 for English System of Units and is 0.01 for 
International System of Units. The second correction term adjusts for 
the effects of atmospheric attenuation due to the difference in the 
sound propagation path length where ALr is the Reference 
takeoff sound propagation path. The third correction term, known as the 
``inverse square'' law, adjusts for the effect of the difference in the 
sound propagation path lengths.
    (ii) Step 2. The corrected values of the SPL(i)r are then 
converted to reference condition PNLT and a correction term calculated 
as follows:

[Delta]1 = PNLT - PNLTM

which represents the correction to be added algebraically to the EPNL 
calculated from the measured data.
    (2) Level flyover flight path. (i) The procedure described in 
paragraph (f)(1) of this section for takeoff paths is also used for the 
level flyover paths, with the values of SPL(i)r relating to 
the flyover sound propagation paths shown in Figure H2 as follows:

SPL(i)r = SPL(i) + C[[alpha](i) - 
          [alpha](i)o]AM + 
          C[alpha](i)o (AM - AMr) + 20 
          log (AM/AMr)

where the lines AM and AMr are the measured and reference 
level flyover sound propagation paths, respectively.
    (ii) The remainder of the procedure is the same for the flyover 
condition as that prescribed in the paragraph (f)(1)(ii) of this section 
regarding takeoff flight path.
    (3) Approach flight path. (i) The procedure described in paragraph 
(f)(1) of this section for takeoff paths is also used for the approach 
paths, with the values of SPL(i)r relating to the approach 
sound propagation paths shown in Figure H3 as follows:

SPL(i)r = SPL(i) + C[[alpha](i) - 
          [alpha](i)o]AN + 
          C[alpha](i)o (AN - ANr) + 20 
          log (AN/ANr)

where the lines AN and ANr are the measured and reference 
approach sound propagation paths, respectively.
    (ii) The remainder of the procedure is the same for the approach 
condition as that prescribed in the paragraph (f)(1)(ii) of this section 
regarding takeoff flight path.
    (4) Sideline microphones. (i) The procedure prescribed in paragraph 
(f)(1) of this section for takeoff paths is also used for the 
propagation to the sideline locations, with the values of 
SPL(i)r relating as follows to the measured sideline sound 
propagation path shown in Figure H3 as follows:

SPL(i)r = SPL(i) + C[[alpha](i) - 
          [alpha](i)o]SX + 
          C[alpha](i)o (SX - SXr) + 20 
          log (SX/SXr)

where S is the sideline measuring station and, based upon the flight 
condition, the helicopter positions, X and Xr, correspond to:


[[Page 1001]]


X = L, and Xr = Lr for takeoff
X = M, and Xr = Mr for flyover
X = N, and Xr = Nr for approach

    (ii) The remainder of the procedure is the same for the sideline 
paths as that prescribed in the paragraph (f)(1)(ii) of this section 
regarding takeoff flight paths.
    (g) Duration corrections. (1) If the measured takeoff and approach 
flight paths do not conform to those prescribed as the corrected and 
reference flight paths, respectively, under section A36.5(d)(2) it will 
be necessary to apply duration corrections to the EPNL values calculated 
from the measured data. Such corrections must be calculated as follows:
    (i) Takeoff flight path. For the takeoff path shown in Figure H1, 
the correction term is calculated using the formula--

[Delta]2 = -7.5 log (AL/ALr) + 10 log (V/
          Vr)

which represents the correction that must be added algebraically to the 
EPNL calculated from the measured data. The lengths AL and 
ALr are the measured and reference takeoff distances from the 
noise measuring station A to the measured and the reference takeoff 
paths, respectively. A negative sign indicates that, for the particular 
case of a duration correction, the EPNL calculated from the measured 
data must be reduced if the measured takeoff path is at greater altitude 
than the reference takeoff path.
    (ii) Level flyover flight paths. For the level flyover flight path, 
the correction term is calculated using the formula--

[Delta]2 = -7.5 log (AM/AMr) + 10 log (V/
          Vr)

where AM is the measured flyover distance from the noise measuring 
station A to the measured flyover path, and AMr is the 
reference distance from station A to the reference flyover path.
    (iii) Approach flight path. For the approach path shown in Figure 
H3, the correction term is calculated using the formula--

[Delta]2 = -7.5 log (AN/ANr) + 10 log (V/
          Vr)

where AN is the measured approach distance from the noise measuring 
station A to the measured approach path, and ANr is the 
reference distance from station A to the reference approach path.
    (iv) Sideline microphones. For the sideline flight path, the 
correction term is calculated using the formula--

[Delta]2 = -7.5 log (SX/SXr) + 10 log (V/
          Vr)

where S is the sideline measuring station and based upon the flight 
condition, the helicopter positions, X and Xr, correspond to:

X = L, and Xr = Lr for takeoff
X = M, and Xr = Mr for flyover
X = N, and Xr = Nr for approach

    (2) The adjustment procedure described in this section shall apply 
to the sideline microphones in the take-off, overflight, and approach 
cases. Although the noise emission is strongly dependent on the 
directivity pattern, variable from one helicopter type to another, the 
propagation angle [thetas] shall be the same for test and reference 
flight paths. The elevation angle [psi] shall not be constrained but 
must be determined and reported. The certification authority shall 
specify the acceptable limitations on [psi]. Corrections to data 
obtained when these limits are exceeded shall be applied using FAA 
approved procedures.

                 part d--noise limits under Sec. 36.805

     Section H36.301 Noise measurement, evaluation, and calculation.

    Compliance with this part of this appendix must be shown with noise 
levels measured, evaluated, and calculated as prescribed under Parts B 
and C of this appendix.

                       Section H36.303 [Reserved]

                      Section H36.305 Noise levels.

    (a) Limits. For compliance with this appendix, the applicant must 
show by flight test that the calculated noise levels of the helicopter, 
at the measuring points described in section H36.305(a) of this 
appendix, do not exceed the following, (with appropriate interpolation 
between weights):
    (1) Stage 1 noise limits for acoustical changes for helicopters are 
as follows:
    (i) For takeoff, flyover, and approach calculated noise levels, the 
noise levels of each Stage 1 helicopter that exceed the Stage 2 noise 
limits plus 2 EPNdB may not, after a change in type design, exceed the 
noise levels created prior to the change in type design.
    (ii) For takeoff, flyover, and approach calculated noise levels, the 
noise levels of each Stage 1 helicopter that do not exceed the Stage 2 
noise limits plus 2 EPNdB may not, after the change in type design, 
exceed the Stage 2 noise limits plus 2 EPNdB.
    (2) Stage 2 noise limits are as follows:
    (i) For takeoff calculated noise levels--109 EPNdB for maximum 
takeoff weights of 176,370 pounds (80,000 kg) or more, reduced by 3.01 
EPNdB per halving of the weight down to 89 EPNdB, after which the limit 
is constant.
    (ii) For flyover calculated noise levels--108 EPNdB for maximum 
weights of 176,370 pounds (80,000 kg) or more, reduced by 3.01 EPNdB per 
halving of the weight down to 88 EPNdB, after which the limit is 
constant.
    (iii) For approach calculated noise levels--110 EPNdB for maximum 
weights of 176,370 pounds (80,000 kg) or more, reduced by 3.01 EPNdB per 
halving of the weight down to 90 EPNdB, after which the limit is 
constant.
    (3) Stage 3 noise limits are as follows:
    (i) For takeoff--For a helicopter having a maximum certificated 
takeoff weight of 176,370 pounds (80,000 kg) or more, the noise limit is 
106 EPNdB, which decreases linearly

[[Page 1002]]

with the logarithm of the helicopter weight (mass) at a rate of 3.0 
EPNdB per halving of the weight (mass) down to 86 EPNdB, after which the 
limit is constant.
    (ii) For flyover--For a helicopter having a maximum certificated 
takeoff weight of 176,370 pounds (80,000 kg) or more, the noise limit is 
104 EPNdB, which decreases linearly with the logarithm of the helicopter 
weight (mass) at a rate of 3.0 EPNdB per halving of the weight (mass) 
down to 84 EPNdB, after which the limit is constant.
    (iii) For approach--For a helicopter having a maximum certificated 
takeoff weight of 176,370 pounds (80,000 kg) or more, the noise limit is 
109 EPNdB, which decreases linearly with the logarithm of the helicopter 
weight (mass) at a rate of 3.0 EPNdB per halving of the weight (mass) 
down to 89 EPNdB, after which the limit is constant.
    (b) Tradeoffs. Except to the extent limited under Sec. 36.11(b) of 
this part, the noise limits prescribed in paragraph (a) of this section 
may be exceeded by one or two of the takeoff, flyover, or approach 
calculated noise levels determined under section H36.203 of this 
appendix if
    (1) The sum of the exceedances is not greater than 4 EPNdB;
    (2) No exceedance is greater than 3 EPNdB; and
    (3) The exceedances are completely offset by reduction in the other 
required calculated noise levels.

[Amdt. 36-14, 53 FR 3541, Feb. 5, 1988; 53 FR 4099, Feb. 11, 1988; 53 FR 
7728, Mar. 10, 1988, as amended by Amdt. 36-54, 67 FR 45237, July 8, 
2002; Amdt. 36-25, 69 FR 31234, June 2, 2004; Amdt. 36-25, 69 FR 41573, 
July 9, 2004; Amdt. 36-30, 79 FR 12045, Mar. 4, 2014]



                  Sec. Appendix I to Part 36 [Reserved]



 Sec. Appendix J to Part 36--Alternative Noise Certification Procedure 
 for Helicopters Under Subpart H Having a Maximum Certificated Takeoff 
                  Weight of Not More Than 7,000 Pounds

                      part a--reference conditions

Sec.

J36.1 General.
J36.3 Reference Test Conditions.
J36.5 [Reserved]

         part b--noise measurement procedure under Sec. 36.801

J36.101 Noise certification test and measurement conditions.
J36.103 [Reserved]
J36.105 Flyover test conditions.
J36.107 [Reserved]
J36.109 Measurement of helicopter noise received on the ground.
J36.111 Reporting requirements.
J36.113 [Reserved]

       part c--noise evaluation and calculation under Sec. 36.803

J36.201 Noise evaluation in SEL.
J36.203 Calculation of noise levels.
J36.205 Detailed data correction procedures.

            part d--noise limits procedure under Sec. 36.805

J36.301 Noise measurement, evaluation, and calculation.
J36.303 [Reserved]
J36.305 Noise limits.

                      part a--reference conditions

                         Section J36.1 General.

    This appendix prescribes the alternative noise certification 
requirements identified under Sec. 36.1 of this part and subpart H of 
this part for helicopters in the primary, normal, transport, and 
restricted categories having maximum certificated takeoff weight of not 
more than 7,000 pounds including:
    (a) The conditions under which an alternative noise certification 
test under subpart H of this part must be conducted and the alternative 
measurement procedure that must be used under Sec. 36.801 of this part 
to measure the helicopter noise during the test;
    (b) The alternative procedures which must be used under Sec. 36.803 
of this part to correct the measured data to the reference conditions 
and to calculate the noise evaluation quantity designated as Sound 
Exposure Level (SEL); and
    (c) The noise limits for which compliance must be shown under Sec. 
36.805 of this part.

                Section J36.3 Reference Test Conditions.

    (a) Meteorological conditions. The following are the noise 
certification reference atmospheric conditions which shall be assumed to 
exist from the surface to the helicopter altitude:
    (1) Sea level pressure of 2116 pounds per square foot (76 
centimeters mercury);
    (2) Ambient temperature of 77 degrees Fahrenheit (25 degrees 
Celsius);
    (3) Relative humidity of 70 percent; and
    (4) Zero wind.
    (b) Reference test site. The reference test site is flat and without 
line-of-sight obstructions across the flight path that encompasses the 
10 dB down points of the A-weighted time history.
    (c) Level flyover reference profile. The reference flyover profile 
is a level flight, 492 feet (150 meters) above ground level as measured 
at the noise measuring station. The reference flyover profile has a 
linear flight track and passes directly over the noise monitoring 
station. Airspeed is stabilized at 0.9VH; 0.9VNE; 
0.45VH + 65 kts (120 km/h); or

[[Page 1003]]

0.45VNE + 65 kts (120 km/h), whichever of the four airspeeds 
is least, and maintained throughout the measured portion of the flyover. 
Rotor speed is stabilized at the maximum normal operating RPM throughout 
the 10 dB-down time interval.
    (1) For noise certification purposes, VH is defined as 
the airspeed in level flight obtained using the minimum specification 
engine power corresponding to maximum continuous power available for sea 
level pressure of 2,116 psf (1,013.25 hPa) at 77 [deg]F (25 [deg]C) 
ambient conditions at the relevant maximum certificated weight. The 
value of VH and VNE used for noise certification 
must be included in the Flight Manual.
    (2) VNE is the never-exceed airspeed.
    (d) The weight of the helicopter shall be the maximum takeoff weight 
at which noise certification is requested.

                        Section J36.5 [Reserved]

         Part B--Noise Measurement Procedure Under Sec. 36.801

  Section J36.101 Noise certification test and measurement conditions.

    (a) General. This section prescribes the conditions under which 
helicopter noise certification tests must be conducted and the 
measurement procedures that must be used to measure helicopter noise 
during each test.
    (b) Test site requirements. (1) The noise measuring station must be 
surrounded by terrain having no excessive sound absorption 
characteristics, such as might be caused by thick, matted, or tall 
grass, shrubs, or wooded areas.
    (2) During the period when the flyover noise measurement is within 
10 dB of the maximum A-weighted sound level, no obstruction that 
significantly influences the sound field from the helicopter may exist 
within a conical space above the noise measuring position (the point on 
the ground vertically below the microphone), the cone is defined by an 
axis normal to the ground and by half-angle 80 degrees from this axis.
    (c) Weather restrictions. The test must be conducted under the 
following atmospheric conditions:
    (1) No rain or other precipitation;
    (2) Ambient air temperature between 36 degrees and 95 degrees 
Fahrenheit (2 degrees and 35 degrees Celsius), inclusively, and relative 
humidity between 20 percent and 95 percent inclusively, except that 
testing may not take place where combinations of temperature and 
relative humidity result in a rate of atmospheric attenuation greater 
than 10 dB per 100 meters (30.5 dB per 1000 ft) in the one-third octave 
band centered at 8 kiloHertz.
    (3) Wind velocity that does not exceed 10 knots (19 km/h) and a 
crosswind component that does not exceed 5 knots (9 km/h). The wind 
shall be determined using a continuous averaging process of no greater 
than 30 seconds;
    (4) Measurements of ambient temperature, relative humidity, wind 
speed, and wind direction must be made between 4 feet (1.2 meters) and 
33 feet (10 meters) above the ground. Unless otherwise approved by the 
FAA, ambient temperature and relative humidity must be measured at the 
same height above the ground.
    (5) No anomalous wind conditions (including turbulence) or other 
anomalous meteorological conditions that will significantly affect the 
noise level of the helicopter when the noise is recorded at the noise 
measuring station; and
    (6) If the measurement site is within 6560 feet (2,000 meters) of a 
fixed meteorological station (such as those found at airports or other 
facilities) the weather measurements reported for temperature, relative 
humidity and wind velocity may be used, if approved by the FAA.
    (d) Helicopter testing procedures. (1) The helicopter testing 
procedures and noise measurements must be conducted and processed in a 
manner which yields the noise evaluation measure designated Sound 
Exposure Level (SEL) as defined in section J36.109(b) of this appendix.
    (2) The helicopter height relative to the noise measurement point 
sufficient to make corrections required under section J36.205 of this 
appendix must be determined by an FAA-approved method that is 
independent of normal flight instrumentation, such as radar tracking, 
theodolite triangulation, laser trajectography, or photographic scaling 
techniques.
    (3) If an applicant demonstrates that the design characteristics of 
the helicopter would prevent flight from being conducted in accordance 
with the reference test conditions prescribed under section J36.3 of 
this appendix, then with FAA approval, the reference test conditions 
used under this appendix may vary from the standard reference test 
conditions, but only to the extent demanded by those design 
characteristics which make compliance with the reference test conditions 
impossible.

                       Section J36.103 [Reserved]

                Section J36.105 Flyover test conditions.

    (a) This section prescribes the flight test conditions and allowable 
random deviations for flyover noise tests conducted under this appendix.
    (b) A test series must consist of at least six flights. The number 
of level flights made with a headwind component must be equal to the 
number of level flights made with a tailwind component over the noise 
measurement station:

[[Page 1004]]

    (1) In level flight and in cruise configuration;
    (2) At a height of 492 feet 50 feet (150 
15 meters) above the ground level at the noise 
measuring station; and
    (3) Within 10 degrees from the zenith.
    (c) Each flyover noise test must be conducted:
    (1) At the reference airspeed specified in section J36.3(c) of this 
appendix, with such airspeed adjusted as necessary to produce the same 
advancing blade tip Mach number as associated with the reference 
conditions;
    (i) Advancing blade tip Mach number (MAT) is defined as 
the ratio of the arithmetic sum of blade tip rotational speed 
(VR) and the helicopter true air speed (VT) over 
the speed of sound (c) at 77 degrees Fahrenheit (1135.6 ft/sec or 346.13 
m/sec) such that MAT=(VR+VT)/c; and
    (ii) The airspeed shall not vary from the adjusted reference 
airspeed by more than 3 knots (5 km/hr) or an equivalent FAA-approved variation from 
the reference advancing blade tip Mach number. The adjusted reference 
airspeed shall be maintained throughout the measured portion of the 
flyover.
    (2) At rotor speed stabilized at the power on maximum normal 
operating rotor RPM (1 percent); and
    (3) With the power stabilized during the period when the measured 
helicopter noise level is within 10 dB of the maximum A-weighted sound 
level (LAMAX).
    (d) The helicopter test weight for each flyover test must be within 
plus 5 percent or minus 10 percent of the maximum takeoff weight for 
which certification under this part is requested.
    (e) The requirements of paragraph (b)(2) of this section 
notwithstanding, flyovers at an FAA-approved lower height may be used 
and the results adjusted to the reference measurement point by an FAA-
approved method if the ambient noise in the test area, measured in 
accordance with the requirements prescribed in section J36.109 of this 
appendix, is found to be within 15 dB(A) of the maximum A-weighted 
helicopter noise level (LAMAX) measured at the noise 
measurement station in accordance with section J36.109 of this appendix.

                       Section J36.107 [Reserved]

 Section J36.109 Measurement of helicopter noise received on the ground.

    (a) General. (1) The helicopter noise measured under this appendix 
for noise certification purposes must be obtained with FAA-approved 
acoustical equipment and measurement practices.
    (2) Paragraph (b) of this section identifies and prescribes the 
specifications for the noise evaluation measurements required under this 
appendix. Paragraphs (c) and (d) of this section prescribe the required 
acoustical equipment specifications. Paragraphs (e) and (f) of this 
section prescribe the calibration and measurement procedures required 
under this appendix.
    (b) Noise unit definition. (1) The value of sound exposure level 
(SEL, or as denoted by symbol, LAE), is defined as the level, 
in decibels, of the time integral of squared `A'-weighted sound pressure 
(PA) over a given time period or event, with reference to the 
square of the standard reference sound pressure (PO) of 20 
micropascals and a reference duration of one second.
    (2) This unit is defined by the expression:
    [GRAPHIC] [TIFF OMITTED] TC28SE91.118
    
Where TO is the reference integration time of one second and 
          (t2-t1) is the integration time 
          interval.

    (3) The integral equation of paragraph (b)(2) of this section can 
also be expressed as:
[GRAPHIC] [TIFF OMITTED] TC28SE91.119

Where LA(t) is the time varying A-weighted sound level.

    (4) The integration time (t2-t1) in practice 
shall not be less than the time interval during which LA(t) 
first rises to within 10 dB(A) of its maximum value (LAMAX) 
and last falls below 10 dB(A) of its maximum value.
    (5) The SEL may be approximated by the following expression:

LAE=LAMAX + A

where A is the duration allowance given by:

A=10 log10 (T)

where T=(t2-t1)/2 and LAMAX is defined 
          as the maximum level, in decibels, of the A-weighted sound 
          pressure (slow response) with reference to the square of the 
          standard reference sound pressure (P0).

    (c) Measurement system. The acoustical measurement system must 
consist of FAA-approved equipment equivalent to the following:
    (1) A microphone system with frequency response that is compatible 
with the measurement and analysis system accuracy prescribed in 
paragraph (d) of this section;
    (2) Tripods or similar microphone mountings that minimize 
interference with the sound energy being measured;
    (3) Recording and reproducing equipment with characteristics, 
frequency response, and dynamic range that are compatible with the 
response and accuracy requirements of paragraph (d) of this section; and

[[Page 1005]]

    (4) The calibration and checking of measurement systems must use the 
procedures described in Section A36.3.9.
    (d) Sensing, recording, and reproducing equipment. (1) The noise 
levels measured from helicopter flyovers under this appendix may be 
determined directly by an integrating sound level meter, or the A-
weighted sound level time history may be written onto a graphic level 
recorder set at ``slow'' response from which the SEL value may be 
determined. With the approval of the FAA, the noise signal may be tape 
recorded for subsequent analysis.
    (i) The SEL values from each flyover test may be directly determined 
from an integrating sound level meter complying with the Standards of 
the International Electrotechnical Commission (IEC) Publication No. 804, 
``Integrating-averaging Sound Level Meters,'' as incorporated by 
reference under Sec. 36.6 of this part, for a Type 1 instrument set at 
``slow'' response.
    (ii) The acoustic signal from the helicopter, along with the 
calibration signals specified under paragraph (e) of this section and 
the background noise signal required under paragraph (f) of this section 
may be recorded on a magnetic tape recorder for subsequent analysis by 
an integrating sound level meter identified in paragraph (d)(1)(i) of 
this section. The record/playback system (including the audio tape) of 
the tape recorder must conform to the requirements prescribed in section 
A36.3.6 of appendix A of this part. The tape recorder shall comply with 
specifications of IEC Publication No. 561, ``Electro-acoustical 
Measuring Equipment for Aircraft Noise Certification,'' as incorporated 
by reference under Sec. 36.6 of this part.
    (iii) The characteristics of the complete system shall comply with 
the recommendations given in IEC Publication No. 651, ``Sound Level 
Meters,'' as incorporated by reference under Sec. 36.6 of this part, 
with regard to the specifications concerning microphone, amplifier, and 
indicating instrument characteristics.
    (iv) The response of the complete system to a sensibly plane 
progressive wave of constant amplitude shall lie within the tolerance 
limits specified in Table IV and Table V for Type 1 instruments in IEC 
Publication No. 651, ``Sound Level Meters,'' as incorporated by 
reference under Sec. 36.6 of this part, for weighting curve ``A'' over 
the frequency range of 45 Hz to 11500 Hz.
    (v) A windscreen must be used with the microphone during each 
measurement of the helicopter flyover noise. Correction for any 
insertion loss produced by the windscreen, as a function of the 
frequency of the acoustic calibration required under paragraph (e) of 
this section, must be applied to the measured data and any correction 
applied must be reported.
    (e) Calibrations. (1) If the helicopter acoustic signal is tape 
recorded for subsequent analysis, the measuring system and components of 
the recording system must be calibrated as prescribed under section 
A36.3.6 of appendix A of this part.
    (2) If the helicopter acoustic signal is directly measured by an 
integrating sound level meter:
    (i) The overall sensitivity of the measuring system shall be checked 
before and after the series of flyover tests and at intervals (not 
exceeding one-hour duration) during the flyover tests using an acoustic 
calibrator using sine wave noise generating a known sound pressure level 
at a known frequency.
    (ii) The performance of equipment in the system will be considered 
satisfactory if, during each day's testing, the variation in the 
calibration value does not exceed 0.5 dB. The SEL data collected during 
the flyover tests shall be adjusted to account for any variation in the 
calibration value.
    (iii) A performance calibration analysis of each piece of 
calibration equipment, including acoustic calibrators, reference 
microphones, and voltage insertion devices, must have been made during 
the six calendar months proceeding the beginning of the helicopter 
flyover series. Each calibration shall be traceable to the National 
Institute of Standards and Technology.
    (f) Noise measurement procedures. (1) The microphone shall be of the 
pressure-sensitive capacitive type designed for nearly uniform grazing 
incidence response. The microphone shall be mounted with the center of 
the sensing element 4 feet (1.2 meters) above the local ground surface 
and shall be oriented for grazing incidence such that the sensing 
element, the diaphragm, is substantially in the plane defined by the 
nominal flight path of the helicopter and the noise measurement station.
    (2) If a tape recorder is used, the frequency response of the 
electrical system must be determined at a level within 10 dB of the 
full-scale reading used during the test, utilizing pink or pseudorandom 
noise.
    (3) The ambient noise, including both acoustical background and 
electrical noise of the measurement systems shall be determined in the 
test area and the system gain set at levels which will be used for 
helicopter noise measurements. If helicopter sound levels do not exceed 
the background sound levels by at least 15 dB(A), flyovers at an FAA-
approved lower height may be used and the results adjusted to the 
reference measurement point by an FAA-approved method.
    (4) If an integrating sound level meter is used to measure the 
helicopter noise, the instrument operator shall monitor the continuous 
A-weighted (slow response) noise levels throughout each flyover to 
ensure that the SEL integration process includes, at minimum, all of the 
noise signal between the

[[Page 1006]]

maximum A-weighted sound level (LAMAX) and the 10 dB down 
points in the flyover time history. The instrument operator shall note 
the actual db(A) levels at the start and stop of the SEL integration 
interval and document these levels along with the value of 
LAMAX and the integration interval (in seconds) for inclusion 
in the noise data submitted as part of the reporting requirements under 
section J36.111(b) of this appendix.

                 Section J36.111 Reporting Requirements.

    (a) General. Data representing physical measurements, and 
corrections to measured data, including corrections to measurements for 
equipment response deviations, must be recorded in permanent form and 
appended to the record. Each correction is subject to FAA approval.
    (b) Data reporting. After the completion of the test the following 
data must be included in the test report furnished to the FAA:
    (1) Measured and corrected sound levels obtained with equipment 
conforming to the standards prescribed in section J36.109 of this 
appendix;
    (2) The type of equipment used for measurement and analysis of all 
acoustic, aircraft performance and flight path, and meteorological data;
    (3) The atmospheric environmental data required to demonstrate 
compliance with this appendix, measured throughout the test period;
    (4) Conditions of local topography, ground cover, or events which 
may interfere with the sound recording;
    (5) The following helicopter information:
    (i) Type, model, and serial numbers, if any, of helicopter, 
engine(s) and rotor(s);
    (ii) Gross dimensions of helicopter, location of engines, rotors, 
type of antitorque system, number of blades for each rotor, and 
reference operating conditions for each engine and rotor;
    (iii) Any modifications of non-standard equipment likely to affect 
the noise characteristics of the helicopter;
    (iv) Maximum takeoff weight for which certification under this 
appendix is requested;
    (v) Aircraft configuration, including landing gear positions;
    (vi) VH or VNE (whichever is less) and the 
adjusted reference airspeed;
    (vii) Aircraft gross weight for each test run;
    (viii) Indicated and true airspeed for each test run;
    (ix) Ground speed, if measured, for each run;
    (x) Helicopter engine performance as determined from aircraft 
instruments and manufacturer's data; and
    (xi) Aircraft flight path above ground level, referenced to the 
elevation of the noise measurement station, in feet, determined by an 
FAA-approved method which is independent of normal flight 
instrumentation, such as radar tracking, theodolite triangulation, laser 
trajectography, or photoscaling techniques; and
    (6) Helicopter position and performance data required to make the 
adjustments prescribed under section J36.205 of this appendix and to 
demonstrate compliance with the performance and position restrictions 
prescribed under section J36.105 of this appendix must be recorded at an 
FAA-approved sampling rate.

                       Section J36.113 [Reserved]

      Part C--Noise Evaluation and Calculations Under Sec. 36.803

                Section J36.201 Noise Evaluation in SEL.

    The noise evaluation measure shall be the sound exposure level (SEL) 
in units of dB(A) as prescribed under section J36.109(b) of this 
appendix. The SEL value for each flyover may be directly determined by 
use of an integrating sound level meter. Specifications for the 
integrating sound level meter and requirements governing the use of such 
instrumentation are prescribed under section J36.109 of this appendix.

              Section J36.203 Calculation of Noise Levels.

    (a) To demonstrate compliance with the noise level limits specified 
under section J36.305 of this appendix, the SEL noise levels from each 
valid flyover, corrected as necessary to reference conditions under 
section J36.205 of this appendix, must be arithmetically averaged to 
obtain a single SEL dB(A) mean value for the flyover series. No 
individual flyover run may be omitted from the averaging process, unless 
otherwise specified or approved by the FAA.
    (b) The minimum sample size acceptable for the helicopter flyover 
certification measurements is six. The number of samples must be large 
enough to establish statistically a 90 percent confidence limit that 
does not exceed 1.5 dB(A).
    (c) All data used and calculations performed under this section, 
including the calculated 90 percent confidence limits, must be 
documented and provided under the reporting requirements of section 
J36.111 of this appendix.

          Section J36.205 Detailed Data Correction Procedures.

    (a) When certification test conditions measured under part B of this 
appendix differ from the reference test conditions prescribed under 
section J36.3 of this appendix, appropriate adjustments shall be made to 
the measured noise data in accordance with the methods set out in 
paragraphs (b) and (c) of

[[Page 1007]]

this section. At minimum, appropriate adjustments shall be made for off-
reference altitude and for the difference between reference airspeed and 
adjusted reference airspeed.
    (b) The adjustment for off-reference altitude may be approximated 
from:

J1=12.5 log10(HT/492) 
          dB;

where J1 is the quantity in decibels that 
          must be algebraically added to the measured SEL noise level to 
          correct for an off-reference flight path, HT is the 
          height, in feet, of the test helicopter when directly over the 
          noise measurement point, and the constant (12.5) accounts for 
          the effects on spherical spreading and duration from the off-
          reference altitude.

    (c) The adjustment for the difference between reference airspeed and 
adjusted reference airspeed is calculated from:

J3=10 log10(VRA/
          VR) dB;

Where J3 is the quantity in decibels that 
          must be algebraically added to the measured SEL noise level to 
          correct for the influence of the adjustment of the reference 
          airspeed on the duration of the measured flyover event as 
          perceived at the noise measurement station, VR is 
          the reference airspeed as prescribed under section J36.3.(c) 
          of this appendix, and VRA is the adjusted reference 
          airspeed as prescribed under section J36.105(c) of this 
          appendix.

    (d) No correction for source noise during the flyover other than the 
variation of source noise accounted for by the adjustment of the 
reference airspeed prescribed for under section J36.105(c) of this 
appendix need be applied.
    (e) No correction for the difference between the reference ground 
speed and the actual ground speed need be applied.
    (f) No correction for off-reference atmospheric attenuation need be 
applied.
    (g) The SEL adjustments must be less than 2.0 dB(A) for differences 
between test and reference flight procedures prescribed under section 
J36.105 of this appendix unless a larger adjustment value is approved by 
the FAA.
    (h) All data used and calculations performed under this section must 
be documented and provided under the reporting requirements specified 
under section J36.111 of this appendix.

            Part D--Noise Limits Procedure Under Sec. 36.805

     Section J36.301 Noise Measurement, Evaluation, and Calculation.

    Compliance with this part of this appendix must be shown with noise 
levels measured, evaluated, and calculated as prescribed under parts B 
and C of this appendix.

                       Section J36.303 [Reserved]

                      Section J36.305 Noise Limits.

    For compliance with this appendix, the calculated noise levels of 
the helicopter, at the measuring point described in section J36.101 of 
this appendix, must be shown to not exceed the following (with 
appropriate interpolation between weights):
    (a) For primary, normal, transport, and restricted category 
helicopters having a maximum certificated takeoff weight of not more 
than 7,000 pounds that are noise tested under this appendix:
    (1) Stage 2 noise limit is constant at 82 decibels SEL for 
helicopters up to 1,737 pounds (787 kg) maximum certificated takeoff 
weight (mass) and increases linearly with the logarithm of the 
helicopter weight at a rate of 3.0 decibels SEL per the doubling of 
weight thereafter. The limit may be calculated by the equation:

LAE(limit) = 82 + 3.0 [log10(MTOW/1737)/
          log10(2)] dB,

where MTOW is the maximum takeoff weight, in pounds, for which 
          certification under this appendix is requested.

    (2) Stage 3 noise limit is constant at 82 decibels SEL for 
helicopters up to 3,125 pounds (1,417 kg) maximum certificated takeoff 
weight (mass) and increases linearly with the logarithm of the 
helicopter weight at a rate of 3.0 decibels SEL per the doubling of 
weight thereafter. The limit may be calculated using the equation:

LAE(limit) = 82 + 3.0 [log10(MTOW/3125)/log10(2)] dB,

where MTOW is the maximum takeoff weight, in pounds.

    (b) The procedures required in this amendment shall be done in 
accordance with the International Electrotechnical Commission IEC 
Publication No. 804, entitled ``Integrating-averaging Sound Level 
Meters,'' First Edition, dated 1985. This incorporation by reference was 
approved by the Director of the Federal Register in accordance with 5 
U.S.C. 552(a) and 1 CFR part 51. Copies may be obtained from the Bureau 
Central de la Commission Electrotechnique Internationale, 1, rue de 
Varembe, Geneva, Switzerland or the American National Standard 
Institute, 1430 Broadway, New York City, New York 10018, or at the 
National Archives and Records Administration (NARA). For information on 
the availability of this material at NARA, call 202-741-6030, or go to: 
http://www.archives.gov/federal--register/code--of--federal--
regulations/ibr--locations.html.

[Doc. No. 26910, 57 FR 42855, Sept. 16, 1992, as amended by Amdt. 36-20, 
57 FR 46243, Oct. 7, 1992; 69 FR 18803, Apr. 9, 2004; Amdt. 36-25, 69 FR 
31234, June 2, 2004; Amdt. 36-30, 79 FR 12045, Mar. 4, 2014]

[[Page 1008]]



  Sec. Appendix K to Part 36--Noise Requirements for Tiltrotors Under 
                                Subpart K

K1 General
K2 Noise Evaluation Measure
K3 Noise Measurement Reference Points
K4 Noise Limits
K5 Trade-offs
K6 Noise Certification Reference Procedures
K7 Test Procedures

                           Section K1 General

    This appendix prescribes noise limits and procedures for measuring 
noise and adjusting the data to standard conditions for tiltrotors as 
specified in Sec. 36.1 of this part.

                   Section K2 Noise Evaluation Measure

    The noise evaluation measure is the effective perceived noise level 
in EPNdB, to be calculated in accordance with section A36.4 of Appendix 
A to this part, except corrections for spectral irregularities must be 
determined using the 50 Hz sound pressure level found in section H36.201 
of Appendix H to this part.

              Section K3 Noise Measurement Reference Points

    The following noise reference points must be used when demonstrating 
tiltrotor compliance with section K6 (Noise Certification Reference 
Procedures) and section K7 (Test Procedures) of this appendix:
    (a) Takeoff reference noise measurement points--
    As shown in Figure K1 below:
    (1) The centerline noise measurement flight path reference point, 
designated A, is located on the ground vertically below the reference 
takeoff flight path. The measurement point is located 1,640 feet (500 m) 
in the horizontal direction of flight from the point Cr where transition 
to climbing flight is initiated, as described in section K6.2 of this 
appendix;
    (2) Two sideline noise measurement points, designated as 
S(starboard) and S(port), are located on the ground perpendicular to and 
symmetrically stationed at 492 feet (150 m) on each side of the takeoff 
reference flight path. The measurement points bisect the centerline 
flight path reference point A.
[GRAPHIC] [TIFF OMITTED] TR08JA13.007

    (b) Flyover reference noise measurement points--
    As shown in Figure K2 below:
    (1) The centerline noise measurement flight path reference point, 
designated A, is located on the ground 492 feet (150 m)

[[Page 1009]]

vertically below the reference flyover flight path. The measurement 
point is defined by the flyover reference procedure in section K6.3 of 
this appendix;
[GRAPHIC] [TIFF OMITTED] TR08JA13.008

    (2) Two sideline noise measurement points, designated as 
S(sideline), are located on the ground perpendicular to and 
symmetrically stationed at 492 feet (150 m) on each side of the flyover 
reference flight path. The measurement points bisect the centerline 
flight path reference point A.
    (c) Approach reference noise measurement points--
    As shown in Figure K3 below:
    (1) The centerline noise measurement flight path reference point, 
designated A, is located on the ground 394 feet (120 m) vertically below 
the reference approach flight path. The measurement point is defined by 
the approach reference procedure in section K6.4 of this appendix. On 
level ground, the measurement point corresponds to a position 3,740 feet 
(1,140 m) from the intersection of the 6.0 degree approach path with the 
ground plane;
    (2) Two sideline noise measurement points, designated as 
S(starboard) and S(port), are located on the ground perpendicular to and 
symmetrically stationed at 492 feet (150 m) on each side of the approach 
reference flight path. The measurement points bisect the centerline 
flight path reference point A.

[[Page 1010]]

[GRAPHIC] [TIFF OMITTED] TR08JA13.009

                         Section K4 Noise Limits

    For a tiltrotor, the maximum noise levels, as determined in 
accordance with the noise evaluation in EPNdB and calculation method 
described in section H36.201 of Appendix H of this part, must not exceed 
the noise limits as follows:
    (a) At the takeoff flight path reference point: For a tiltrotor 
having a maximum certificated takeoff weight (mass) of 176,370 pounds 
(80,000 kg) or more, in VTOL/Conversion mode, 109 EPNdB, decreasing 
linearly with the logarithm of the tiltrotor weight (mass) at a rate of 
3.0 EPNdB per halving of weight (mass) down to 89 EPNdB, after which the 
limit is constant. Figure K4 illustrates the takeoff noise limit as a 
solid line.
    (b) At the Flyover path reference point: For a tiltrotor having a 
maximum certificated takeoff weight (mass) of 176,370 pounds (80,000 kg) 
or more, in VTOL/Conversion mode, 108 EPNdB, decreasing linearly with 
the logarithm of the tiltrotor weight (mass) at a rate of 3.0 EPNdB per 
halving of weight (mass) down to 88 EPNdB, after which the limit is 
constant. Figure K4 illustrates the flyover noise limit as a dashed 
line.
    (c) At the approach flight path reference point: For a tiltrotor 
having a maximum certificated takeoff weight (mass) of 176,370 pounds 
(80,000 kg) or more, in VTOL/Conversion mode, 110 EPNdB, decreasing 
linearly with the logarithm of the tiltrotors weight (mass) at a rate of 
3.0 EPNdB per halving of weight (mass) down to 90 EPNdB, after which the 
limit is constant. Figure K4 illustrates the approach noise limit as a 
dash-dot line.

[[Page 1011]]

[GRAPHIC] [TIFF OMITTED] TR08JA13.010

                          Section K5 Trade-Offs

    If the noise evaluation measurement exceeds the noise limits 
described in K4 of this appendix at one or two measurement points:
    (a) The sum of excesses must not be greater than 4 EPNdB;
    (b) The excess at any single point must not be greater than 3 EPNdB; 
and
    (c) Any excess must be offset by the remaining noise margin at the 
other point or points.

           Section K6 Noise Certification Reference Procedures

    K6.1 General Conditions
    (a)-(b) [Reserved]
    (c) The takeoff, flyover and approach reference procedures must be 
established in accordance with sections K6.2, K6.3 and K6.4 of this 
appendix, except as specified in section K6.1(d) of this appendix.
    (d) If the design characteristics of the tiltrotor prevent test 
flights from being conducted in accordance with section K6.2, K6.3 or 
K6.4 of this appendix, the applicant must revise the test procedures and 
resubmit the procedures for approval.
    (e) The following reference atmospheric conditions must be used to 
establish the reference procedures:
    (1) Sea level atmospheric pressure of 2,116 pounds per square foot 
(1,013.25 hPa);
    (2) Ambient air temperature of 77 [deg]Fahrenheit (25 [deg]Celsius, 
i.e. ISA + 10 [deg]C);
    (3) Relative humidity of 70 percent; and
    (4) Zero wind.
    (f) For tests conducted in accordance with sections K6.2, K6.3, and 
K6.4 of this appendix, use the maximum normal operating RPM

[[Page 1012]]

corresponding to the airworthiness limit imposed by the manufacturer. 
For configurations for which the rotor speed automatically links with 
the flight condition, use the maximum normal operating rotor speed 
corresponding with the reference flight condition. For configurations 
for which the rotor speed can change by pilot action, use the highest 
normal rotor speed specified in the flight manual limitation section for 
the reference conditions.
    K6.2 Takeoff Reference Procedure. The takeoff reference flight 
procedure is as follows:
    (a) A constant takeoff configuration must be maintained, including 
the nacelle angle selected by the applicant;
    (b) The tiltrotor power must be stabilized at the maximum takeoff 
power corresponding to the minimum installed engine(s) specification 
power available for the reference ambient conditions or gearbox torque 
limit, whichever is lower. The tiltrotor power must also be stabilized 
along a path starting from a point located 1,640 feet (500 m) before the 
flight path reference point, at 65 ft (20 m) above ground level;
    (c) The nacelle angle and the corresponding best rate of climb 
speed, or the lowest approved speed for the climb after takeoff, 
whichever is the greater, must be maintained throughout the takeoff 
reference procedure;
    (d) The rotor speed must be stabilized at the maximum normal 
operating RPM certificated for takeoff;
    (e) The weight (mass) of the tiltrotors must be the maximum takeoff 
weight (mass) as requested for noise certification; and
    (f) The reference takeoff flight profile is a straight line segment 
inclined from the starting point 1,640 feet (500 m) before to the center 
noise measurement point and 65 ft (20 m) above ground level at an angle 
defined by best rate of climb and the speed corresponding to the 
selected nacelle angle and for minimum specification engine performance.
    K6.3 Flyover Reference Procedure. The flyover reference flight 
procedure is as follows:
    (a) The tiltrotor must be stabilized for level flight along the 
centerline flyover flight path and over the noise measurement reference 
point at an altitude of 492 ft (150 m) above ground level;
    (b) A constant flyover configuration selected by the applicant must 
be maintained;
    (c) The weight (mass) of the tiltrotor must be the maximum takeoff 
weight (mass) as requested for noise certification;
    (d) In the VTOL/Conversion mode:
    (1) The nacelle angle must be at the authorized fixed operation 
point that is closest to the shallow nacelle angle certificated for zero 
airspeed;
    (2) The airspeed must be 0.9VCON and
    (3) The rotor speed must be stabilized at the maximum normal 
operating RPM certificated for level flight.
    K6.4 Approach Reference Procedure. The approach reference procedure 
is as follows:
    (a) The tiltrotor must be stabilized to follow a 6.0 degree approach 
path;
    (b) An approved airworthiness configuration in which maximum noise 
occurs must be maintained;
    (1) An airspeed equal to the best rate of climb speed corresponding 
to the nacelle angle, or the lowest approved airspeed for the approach, 
whichever is greater, must be stabilized and maintained; and
    (2) The tiltrotor power during the approach must be stabilized over 
the flight path reference point, and continue as if landing;
    (c) The rotor speed must be stabilized at the maximum normal 
operating RPM certificated for approach;
    (d) The constant approach configuration used in airworthiness 
certification tests, with the landing gear extended, must be maintained; 
and
    (e) The weight (mass) of the tiltrotor at landing must be the 
maximum landing weight (mass) as requested for noise certification.

                       Section K7 Test Procedures

    K7.1 [Reserved]
    K7.2 The test procedures and noise measurements must be conducted 
and processed to yield the noise evaluation measure designated in 
section K2 of this appendix.
    K7.3 If either the test conditions or test procedures do not comply 
to the applicable noise certification reference conditions or procedures 
prescribed by this part, the applicant must apply the correction methods 
described in section H36.205 of Appendix H of this part to the acoustic 
test data measured.
    K7.4 Adjustments for differences between test and reference flight 
procedures must not exceed:
    (a) For takeoff: 4.0 EPNdB, of which the arithmetic sum of delta 1 
and the term -7.5 log (QK/QrKr) from delta 2 must not in total exceed 
2.0 EPNdB;
    (b) For flyover or approach: 2.0 EPNdB.
    K7.5 The average rotor RPM must not vary from the normal maximum 
operating RPM by more than 1.0 percent throughout 
the 10 dB-down time interval.
    K7.6 The tiltrotor airspeed must not vary from the reference 
airspeed appropriate to the flight demonstration by more than 5 kts (9 km/h) throughout the 10 
dB-down time interval.
    K7.7 The number of level flyovers made with a head wind component 
must be equal to the number of level flyovers made with a tail wind 
component.
    K7.8 The tiltrotor must operate between 10 
degrees from the vertical or between 65 feet 
(20 m) lateral deviation tolerance,

[[Page 1013]]

whichever is greater, above the reference track and throughout the 10 
dB-down time interval.
    K7.9 The tiltrotor altitude must not vary during each flyover by 
more than 30 ft (9 m) from 
the reference altitude throughout the 10 dB-down time interval.
    K7.10 During the approach procedure, the tiltrotor must establish a 
stabilized constant speed approach and fly between approach angles of 
5.5 degrees and 6.5 degrees throughout the 10 dB-down time interval.
    K7.11 During all test procedures, the tiltrotor weight (mass) must 
not be less than 90 percent and not more than 105 percent of the maximum 
certificated weight (mass). For each of the test procedures, complete at 
least one test at or above this maximum certificated weight (mass).
    K7.12 A tiltrotor capable of carrying external loads or external 
equipment must be noise certificated without such loads or equipment 
fitted
    K7.13 The value of VCON used for noise certification must 
be included in the approved Flight Manual.

[78 FR 1139, Jan. 8, 2013]



PART 39_AIRWORTHINESS DIRECTIVES--Table of Contents



Sec.
39.1 Purpose of this regulation.
39.3 Definition of airworthiness directives.
39.5 When does FAA issue airworthiness directives?
39.7 What is the legal effect of failing to comply with an airworthiness 
          directive?
39.9 What if I operate an aircraft or use a product that does not meet 
          the requirements of an airworthiness directive?
39.11 What actions do airworthiness directives require?
39.13 Are airworthiness directives part of the Code of Federal 
          Regulations?
39.15 Does an airworthiness directive apply if the product has been 
          changed?
39.17 What must I do if a change in a product affects my ability to 
          accomplish the actions required in an airworthiness directive?
39.19 May I address the unsafe condition in a way other than that set 
          out in the airworthiness directive?
39.21 Where can I get information about FAA-approved alternative methods 
          of compliance?
39.23 May I fly my aircraft to a repair facility to do the work required 
          by an airworthiness directive?
39.25 How do I get a special flight permit?
39.27 What do I do if the airworthiness directive conflicts with the 
          service document on which it is based?

    Authority: 49 U.S.C. 106(g), 40113, 44701.

    Source: Docket No. FAA-2000-8460, 67 FR 48003, July 22, 2002, unless 
otherwise noted.



Sec. 39.1  Purpose of this regulation.

    The regulations in this part provide a legal framework for FAA's 
system of Airworthiness Directives.



Sec. 39.3  Definition of airworthiness directives.

    FAA's airworthiness directives are legally enforceable rules that 
apply to the following products: aircraft, aircraft engines, propellers, 
and appliances.



Sec. 39.5  When does FAA issue airworthiness directives?

    FAA issues an airworthiness directive addressing a product when we 
find that:
    (a) An unsafe condition exists in the product; and
    (b) The condition is likely to exist or develop in other products of 
the same type design.



Sec. 39.7  What is the legal effect of failing to comply with an 
airworthiness directive?

    Anyone who operates a product that does not meet the requirements of 
an applicable airworthiness directive is in violation of this section.



Sec. 39.9  What if I operate an aircraft or use a product that does not
meet the requirements of an airworthiness directive?

    If the requirements of an airworthiness directive have not been met, 
you violate Sec. 39.7 each time you operate the aircraft or use the 
product.



Sec. 39.11  What actions do airworthiness directives require?

    Airworthiness directives specify inspections you must carry out, 
conditions and limitations you must comply with, and any actions you 
must take to resolve an unsafe condition.



Sec. 39.13  Are airworthiness directives part of the Code of Federal
Regulations?

    Yes, airworthiness directives are part of the Code of Federal 
Regulations, but

[[Page 1014]]

they are not codified in the annual edition. FAA publishes airworthiness 
directives in full in the Federal Register as amendments to Sec. 39.13.

    Editorial Note: For a complete list of citations to airworthiness 
directives published in the Federal Register, consult the following 
publications: For airworthiness directives published in the Federal 
Register since 2001, see the entries for 14 CFR 39.13 in the List of CFR 
Sections Affected, which appears in the ``Finding Aids'' section of the 
printed volume and at www.fdsys.gov. For citations to prior amendments, 
see the entries for 14 CFR 39.13 in the separate publications List of 
CFR Sections Affected, 1973-1985, List of CFR Sections Affected, 1964-
1972, and List of CFR Sections Affected, 1986-2000, and the entries for 
14 CFR 507.10 in the List of Sections Affected, 1949-1963. See also the 
annual editions of the Federal Register Index for subject matter 
references and citations to FAA airworthiness directives.



Sec. 39.15  Does an airworthiness directive apply if the product has
been changed?

    Yes, an airworthiness directive applies to each product identified 
in the airworthiness directive, even if an individual product has been 
changed by modifying, altering, or repairing it in the area addressed by 
the airworthiness directive.



Sec. 39.17  What must I do if a change in a product affects my ability
to accomplish the actions required in an airworthiness directive?

    If a change in a product affects your ability to accomplish the 
actions required by the airworthiness directive in any way, you must 
request FAA approval of an alternative method of compliance. Unless you 
can show the change eliminated the unsafe condition, your request should 
include the specific actions that you propose to address the unsafe 
condition. Submit your request in the manner described in Sec. 39.19.



Sec. 39.19  May I address the unsafe condition in a way other than 
that set out in the airworthiness directive?

    Yes, anyone may propose to FAA an alternative method of compliance 
or a change in the compliance time, if the proposal provides an 
acceptable level of safety. Unless FAA authorizes otherwise, send your 
proposal to your principal inspector. Include the specific actions you 
are proposing to address the unsafe condition. The principal inspector 
may add comments and will send your request to the manager of the office 
identified in the airworthiness directive (manager). You may send a copy 
to the manager at the same time you send it to the principal inspector. 
If you do not have a principal inspector send your proposal directly to 
the manager. You may use the alternative you propose only if the manager 
approves it.



Sec. 39.21  Where can I get information about FAA-approved alternative
methods of compliance?

    Each airworthiness directive identifies the office responsible for 
approving alternative methods of compliance. That office can provide 
information about alternatives it has already approved.



Sec. 39.23  May I fly my aircraft to a repair facility to do the work
required by an airworthiness directive?

    Yes, the operations specifications giving some operators authority 
to operate include a provision that allow them to fly their aircraft to 
a repair facility to do the work required by an airworthiness directive. 
If you do not have this authority, the local Flight Standards District 
Office of FAA may issue you a special flight permit unless the 
airworthiness directive states otherwise. To ensure aviation safety, FAA 
may add special requirements for operating your aircraft to a place 
where the repairs or modifications can be accomplished. FAA may also 
decline to issue a special flight permit in particular cases if we 
determine you cannot move the aircraft safely.



Sec. 39.25  How do I get a special flight permit?

    Apply to FAA for a special flight permit following the procedures in 
14 CFR 21.199.

[[Page 1015]]



Sec. 39.27  What do I do if the airworthiness directive conflicts with
the service document on which it is based?

    In some cases an airworthiness directive incorporates by reference a 
manufacturer's service document. In these cases, the service document 
becomes part of the airworthiness directive. In some cases the 
directions in the service document may be modified by the airworthiness 
directive. If there is a conflict between the service document and the 
airworthiness directive, you must follow the requirements of the 
airworthiness directive.



PART 43_MAINTENANCE, PREVENTIVE MAINTENANCE, REBUILDING, AND ALTERATION
--Table of Contents



Sec.
43.1 Applicability.
43.2 Records of overhaul and rebuilding.
43.3 Persons authorized to perform maintenance, preventive maintenance, 
          rebuilding, and alterations.
43.5 Approval for return to service after maintenance, preventive 
          maintenance, rebuilding, or alteration.
43.7 Persons authorized to approve aircraft, airframes, aircraft 
          engines, propellers, appliances, or component parts for return 
          to service after maintenance, preventive maintenance, 
          rebuilding, or alteration.
43.9 Content, form, and disposition of maintenance, preventive 
          maintenance, rebuilding, and alteration records (except 
          inspections performed in accordance with part 91, part 123, 
          part 125, Sec. 135.411(a)(1), and Sec. 135.419 of this 
          chapter).
43.10 Disposition of life-limited aircraft parts.
43.11 Content, form, and disposition of records for inspections 
          conducted under parts 91 and 125 and Sec. Sec. 135.411(a)(1) 
          and 135.419 of this chapter.
43.12 Maintenance records: Falsification, reproduction, or alteration.
43.13 Performance rules (general).
43.15 Additional performance rules for inspections.
43.16 Airworthiness Limitations.
43.17 Maintenance, preventive maintenance, and alterations performed on 
          U.S. aeronautical products by certain Canadian persons.

Appendix A to Part 43--Major Alterations, Major Repairs, and Preventive 
          Maintenance
Appendix B to Part 43--Recording of Major Repairs and Major Alterations
Appendix C to Part 43 [Reserved]
Appendix D to Part 43--Scope and Detail of Items (as Applicable to the 
          Particular Aircraft) To Be Included in Annual and 100-Hour 
          Inspections
Appendix E to Part 43--Altimeter System Test and Inspection
Appendix F to Part 43--ATC Transponder Tests and Inspections

    Authority: 49 U.S.C. 106(g), 40113, 44701, 44703, 44705, 44707, 
44711, 44713, 44717, 44725.

    Source: Docket No. 1993, 29 FR 5451, Apr. 23, 1964, unless otherwise 
noted.

    Editorial Note: For miscellaneous technical amendments to this part 
43, see Amdt. 43-3, 31 FR 3336, Mar. 3, 1966, and Amdt. 43-6, 31 FR 
9211, July 6, 1966.



Sec. 43.1  Applicability.

    (a) Except as provided in paragraphs (b) and (d) of this section, 
this part prescribes rules governing the maintenance, preventive 
maintenance, rebuilding, and alteration of any--
    (1) Aircraft having a U.S. airworthiness certificate;
    (2) Foreign-registered civil aircraft used in common carriage or 
carriage of mail under the provisions of Part 121 or 135 of this 
chapter; and
    (3) Airframe, aircraft engines, propellers, appliances, and 
component parts of such aircraft.
    (b) This part does not apply to--
    (1) Any aircraft for which the FAA has issued an experimental 
certificate, unless the FAA has previously issued a different kind of 
airworthiness certificate for that aircraft; or
    (2) Any aircraft for which the FAA has issued an experimental 
certificate under the provisions of Sec. 21.191 (i)(3) of this chapter, 
and the aircraft was previously issued a special airworthiness 
certificate in the light-sport category under the provisions of Sec. 
21.190 of this chapter.
    (c) This part applies to all life-limited parts that are removed 
from a type certificated product, segregated, or controlled as provided 
in Sec. 43.10.
    (d) This part applies to any aircraft issued a special airworthiness 
certificate in the light-sport category except:
    (1) The repair or alteration form specified in Sec. Sec. 43.5(b) 
and 43.9(d) is not required to be completed for products not produced 
under an FAA approval;

[[Page 1016]]

    (2) Major repairs and major alterations for products not produced 
under an FAA approval are not required to be recorded in accordance with 
appendix B of this part; and
    (3) The listing of major alterations and major repairs specified in 
paragraphs (a) and (b) of appendix A of this part is not applicable to 
products not produced under an FAA approval.

[Doc. No. 1993, 29 FR 5451, Apr. 23, 1964, as amended by Amdt. 43-23, 47 
FR 41084, Sept. 16, 1982; Amdt. 43-37, 66 FR 21066, Apr. 27, 2001; Amdt. 
43-38, 67 FR 2109, Jan. 15, 2002; Amdt. 43-39, 69 FR 44863, July 27, 
2004; Amdt. 43-44, 75 FR 5219, Feb. 1, 2010]



Sec. 43.2  Records of overhaul and rebuilding.

    (a) No person may describe in any required maintenance entry or form 
an aircraft, airframe, aircraft engine, propeller, appliance, or 
component part as being overhauled unless--
    (1) Using methods, techniques, and practices acceptable to the 
Administrator, it has been disassembled, cleaned, inspected, repaired as 
necessary, and reassembled; and
    (2) It has been tested in accordance with approved standards and 
technical data, or in accordance with current standards and technical 
data acceptable to the Administrator, which have been developed and 
documented by the holder of the type certificate, supplemental type 
certificate, or a material, part, process, or appliance approval under 
part 21 of this chapter.
    (b) No person may describe in any required maintenance entry or form 
an aircraft, airframe, aircraft engine, propeller, appliance, or 
component part as being rebuilt unless it has been disassembled, 
cleaned, inspected, repaired as necessary, reassembled, and tested to 
the same tolerances and limits as a new item, using either new parts or 
used parts that either conform to new part tolerances and limits or to 
approved oversized or undersized dimensions.

[Amdt. 43-23, 47 FR 41084, Sept. 16, 1982, as amended by Amdt. 43-43, 74 
FR 53394, Oct. 16, 2009]



Sec. 43.3  Persons authorized to perform maintenance, preventive
maintenance, rebuilding, and alterations.

    (a) Except as provided in this section and Sec. 43.17, no person 
may maintain, rebuild, alter, or perform preventive maintenance on an 
aircraft, airframe, aircraft engine, propeller, appliance, or component 
part to which this part applies. Those items, the performance of which 
is a major alteration, a major repair, or preventive maintenance, are 
listed in appendix A.
    (b) The holder of a mechanic certificate may perform maintenance, 
preventive maintenance, and alterations as provided in Part 65 of this 
chapter.
    (c) The holder of a repairman certificate may perform maintenance, 
preventive maintenance, and alterations as provided in part 65 of this 
chapter.
    (d) A person working under the supervision of a holder of a mechanic 
or repairman certificate may perform the maintenance, preventive 
maintenance, and alterations that his supervisor is authorized to 
perform, if the supervisor personally observes the work being done to 
the extent necessary to ensure that it is being done properly and if the 
supervisor is readily available, in person, for consultation. However, 
this paragraph does not authorize the performance of any inspection 
required by Part 91 or Part 125 of this chapter or any inspection 
performed after a major repair or alteration.
    (e) The holder of a repair station certificate may perform 
maintenance, preventive maintenance, and alterations as provided in Part 
145 of this chapter.
    (f) The holder of an air carrier operating certificate or an 
operating certificate issued under Part 121 or 135, may perform 
maintenance, preventive maintenance, and alterations as provided in Part 
121 or 135.
    (g) Except for holders of a sport pilot certificate, the holder of a 
pilot certificate issued under part 61 may perform preventive 
maintenance on any aircraft owned or operated by that pilot which is not 
used under part 121, 129, or 135 of this chapter. The holder of a sport 
pilot certificate may perform preventive maintenance on an aircraft 
owned or operated by that pilot and issued a special airworthiness 
certificate in the light-sport category.

[[Page 1017]]

    (h) Notwithstanding the provisions of paragraph (g) of this section, 
the Administrator may approve a certificate holder under Part 135 of 
this chapter, operating rotorcraft in a remote area, to allow a pilot to 
perform specific preventive maintenance items provided--
    (1) The items of preventive maintenance are a result of a known or 
suspected mechanical difficulty or malfunction that occurred en route to 
or in a remote area;
    (2) The pilot has satisfactorily completed an approved training 
program and is authorized in writing by the certificate holder for each 
item of preventive maintenance that the pilot is authorized to perform;
    (3) There is no certificated mechanic available to perform 
preventive maintenance;
    (4) The certificate holder has procedures to evaluate the 
accomplishment of a preventive maintenance item that requires a decision 
concerning the airworthiness of the rotorcraft; and
    (5) The items of preventive maintenance authorized by this section 
are those listed in paragraph (c) of appendix A of this part.
    (i) Notwithstanding the provisions of paragraph (g) of this section, 
in accordance with an approval issued to the holder of a certificate 
issued under part 135 of this chapter, a pilot of an aircraft type-
certificated for 9 or fewer passenger seats, excluding any pilot seat, 
may perform the removal and reinstallation of approved aircraft cabin 
seats, approved cabin-mounted stretchers, and when no tools are 
required, approved cabin-mounted medical oxygen bottles, provided--
    (1) The pilot has satisfactorily completed an approved training 
program and is authorized in writing by the certificate holder to 
perform each task; and
    (2) The certificate holder has written procedures available to the 
pilot to evaluate the accomplishment of the task.
    (j) A manufacturer may--
    (1) Rebuild or alter any aircraft, aircraft engine, propeller, or 
appliance manufactured by him under a type or production certificate;
    (2) Rebuild or alter any appliance or part of aircraft, aircraft 
engines, propellers, or appliances manufactured by him under a Technical 
Standard Order Authorization, an FAA-Parts Manufacturer Approval, or 
Product and Process Specification issued by the Administrator; and
    (3) Perform any inspection required by part 91 or part 125 of this 
chapter on aircraft it manufactured under a type certificate, or 
currently manufactures under a production certificate.
    (k) Updates of databases in installed avionics meeting the 
conditions of this paragraph are not considered maintenance and may be 
performed by pilots provided:
    (1) The database upload is:
    (i) Initiated from the flight deck;
    (ii) Performed without disassembling the avionics unit; and
    (iii) Performed without the use of tools and/or special equipment.
    (2) The pilot must comply with the certificate holder's procedures 
or the manufacturer's instructions.
    (3) The holder of operating certificates must make available written 
procedures consistent with manufacturer's instructions to the pilot that 
describe how to:
    (i) Perform the database update; and
    (ii) Determine the status of the data upload.

[Doc. No. 1993, 29 FR 5451, Apr. 23, 1964, as amended by Amdt. 43-4, 31 
FR 5249, Apr. 1, 1966; Amdt. 43-23, 47 FR 41084, Sept. 16, 1982; Amdt. 
43-25, 51 FR 40702, Nov. 7, 1986; Amdt. 43-36, 61 FR 19501, May 1, 1996; 
Amdt. 43-37, 66 FR 21066, Apr. 27, 2001; Amdt. 43-39, 69 FR 44863, July 
27, 2004; Amdt. 43-43, 74 FR 53394, Oct. 16, 2009; Amdt. 43-45, 77 FR 
71096, Nov. 29, 2012]



Sec. 43.5  Approval for return to service after maintenance, preventive
maintenance, rebuilding, or alteration.

    No person may approve for return to service any aircraft, airframe, 
aircraft engine, propeller, or appliance, that has undergone 
maintenance, preventive maintenance, rebuilding, or alteration unless--
    (a) The maintenance record entry required by Sec. 43.9 or Sec. 
43.11, as appropriate, has been made;
    (b) The repair or alteration form authorized by or furnished by the 
Administrator has been executed in a manner prescribed by the 
Administrator; and

[[Page 1018]]

    (c) If a repair or an alteration results in any change in the 
aircraft operating limitations or flight data contained in the approved 
aircraft flight manual, those operating limitations or flight data are 
appropriately revised and set forth as prescribed in Sec. 91.9 of this 
chapter.

[Doc. No. 1993, 29 FR 5451, Apr. 23, 1964, as amended by Amdt. 43-23, 47 
FR 41084, Sept. 16, 1982; Amdt. 43-31, 54 FR 34330, Aug. 18, 1989]



Sec. 43.7  Persons authorized to approve aircraft, airframes, aircraft
engines, propellers, appliances, or component parts for return to 
service after maintenance, preventive maintenance, rebuilding, or 
          alteration.

    (a) Except as provided in this section and Sec. 43.17, no person, 
other than the Administrator, may approve an aircraft, airframe, 
aircraft engine, propeller, appliance, or component part for return to 
service after it has undergone maintenance, preventive maintenance, 
rebuilding, or alteration.
    (b) The holder of a mechanic certificate or an inspection 
authorization may approve an aircraft, airframe, aircraft engine, 
propeller, appliance, or component part for return to service as 
provided in Part 65 of this chapter.
    (c) The holder of a repair station certificate may approve an 
aircraft, airframe, aircraft engine, propeller, appliance, or component 
part for return to service as provided in Part 145 of this chapter.
    (d) A manufacturer may approve for return to service any aircraft, 
airframe, aircraft engine, propeller, appliance, or component part which 
that manufacturer has worked on under Sec. 43.3(j). However, except for 
minor alterations, the work must have been done in accordance with 
technical data approved by the Administrator.
    (e) The holder of an air carrier operating certificate or an 
operating certificate issued under Part 121 or 135, may approve an 
aircraft, airframe, aircraft engine, propeller, appliance, or component 
part for return to service as provided in Part 121 or 135 of this 
chapter, as applicable.
    (f) A person holding at least a private pilot certificate may 
approve an aircraft for return to service after performing preventive 
maintenance under the provisions of Sec. 43.3(g).
    (g) The holder of a repairman certificate (light-sport aircraft) 
with a maintenance rating may approve an aircraft issued a special 
airworthiness certificate in light-sport category for return to service, 
as provided in part 65 of this chapter.
    (h) The holder of at least a sport pilot certificate may approve an 
aircraft owned or operated by that pilot and issued a special 
airworthiness certificate in the light-sport category for return to 
service after performing preventive maintenance under the provisions of 
Sec. 43.3(g).

[Amdt. 43-23, 47 FR 41084, Sept. 16, 1982, as amended by Amdt. 43-36, 61 
FR 19501, May 1, 1996; Amdt. 43-37, 66 FR 21066, Apr. 27, 2001; Amdt. 
43-39, 69 FR 44863, July 27, 2004]



Sec. 43.9  Content, form, and disposition of maintenance, preventive
maintenance, rebuilding, and alteration records (except inspections 
performed in accordance with part 91, part 125, Sec. 135.411(a)(1), 
and  Sec. 135.419 of this chapter).

    (a) Maintenance record entries. Except as provided in paragraphs (b) 
and (c) of this section, each person who maintains, performs preventive 
maintenance, rebuilds, or alters an aircraft, airframe, aircraft engine, 
propeller, appliance, or component part shall make an entry in the 
maintenance record of that equipment containing the following 
information:
    (1) A description (or reference to data acceptable to the 
Administrator) of work performed.
    (2) The date of completion of the work performed.
    (3) The name of the person performing the work if other than the 
person specified in paragraph (a)(4) of this section.
    (4) If the work performed on the aircraft, airframe, aircraft 
engine, propeller, appliance, or component part has been performed 
satisfactorily, the signature, certificate number, and kind of 
certificate held by the person approving the work. The signature 
constitutes the approval for return to service only for the work 
performed.
    (b) Each holder of an air carrier operating certificate or an 
operating certificate issued under Part 121 or 135,

[[Page 1019]]

that is required by its approved operations specifications to provide 
for a continuous airworthiness maintenance program, shall make a record 
of the maintenance, preventive maintenance, rebuilding, and alteration, 
on aircraft, airframes, aircraft engines, propellers, appliances, or 
component parts which it operates in accordance with the applicable 
provisions of Part 121 or 135 of this chapter, as appropriate.
    (c) This section does not apply to persons performing inspections in 
accordance with Part 91, 125, Sec. 135.411(a)(1), or Sec. 135.419 of 
this chapter.
    (d) In addition to the entry required by paragraph (a) of this 
section, major repairs and major alterations shall be entered on a form, 
and the form disposed of, in the manner prescribed in appendix B, by the 
person performing the work.

[Amdt. 43-23, 47 FR 41085, Sept. 16, 1982, as amended by Amdt. 43-37, 66 
FR 21066, Apr. 27, 2001; Amdt. 43-39, 69 FR 44863, July 27, 2004]



Sec. 43.10  Disposition of life-limited aircraft parts.

    (a) Definitions used in this section. For the purposes of this 
section the following definitions apply.
    Life-limited part means any part for which a mandatory replacement 
limit is specified in the type design, the Instructions for Continued 
Airworthiness, or the maintenance manual.
    Life status means the accumulated cycles, hours, or any other 
mandatory replacement limit of a life-limited part.
    (b) Temporary removal of parts from type-certificated products. When 
a life-limited part is temporarily removed and reinstalled for the 
purpose of performing maintenance, no disposition under paragraph (c) of 
this section is required if--
    (1) The life status of the part has not changed;
    (2) The removal and reinstallation is performed on the same serial 
numbered product; and
    (3) That product does not accumulate time in service while the part 
is removed.
    (c) Disposition of parts removed from type-certificated products. 
Except as provided in paragraph (b) of this section, after April 15, 
2002 each person who removes a life-limited part from a type-
certificated product must ensure that the part is controlled using one 
of the methods in this paragraph. The method must deter the installation 
of the part after it has reached its life limit. Acceptable methods 
include:
    (1) Record keeping system. The part may be controlled using a record 
keeping system that substantiates the part number, serial number, and 
current life status of the part. Each time the part is removed from a 
type certificated product, the record must be updated with the current 
life status. This system may include electronic, paper, or other means 
of record keeping.
    (2) Tag or record attached to part. A tag or other record may be 
attached to the part. The tag or record must include the part number, 
serial number, and current life status of the part. Each time the part 
is removed from a type certificated product, either a new tag or record 
must be created, or the existing tag or record must be updated with the 
current life status.
    (3) Non-permanent marking. The part may be legibly marked using a 
non-permanent method showing its current life status. The life status 
must be updated each time the part is removed from a type certificated 
product, or if the mark is removed, another method in this section may 
be used. The mark must be accomplished in accordance with the 
instructions under Sec. 45.16 of this chapter in order to maintain the 
integrity of the part.
    (4) Permanent marking. The part may be legibly marked using a 
permanent method showing its current life status. The life status must 
be updated each time the part is removed from a type certificated 
product. Unless the part is permanently removed from use on type 
certificated products, this permanent mark must be accomplished in 
accordance with the instructions under Sec. 45.16 of this chapter in 
order to maintain the integrity of the part.
    (5) Segregation. The part may be segregated using methods that deter 
its installation on a type-certificated product. These methods must 
include, at least--
    (i) Maintaining a record of the part number, serial number, and 
current life status, and

[[Page 1020]]

    (ii) Ensuring the part is physically stored separately from parts 
that are currently eligible for installation.
    (6) Mutilation. The part may be mutilated to deter its installation 
in a type certificated product. The mutilation must render the part 
beyond repair and incapable of being reworked to appear to be airworthy.
    (7) Other methods. Any other method approved or accepted by the FAA.
    (d) Transfer of life-limited parts. Each person who removes a life-
limited part from a type certificated product and later sells or 
otherwise transfers that part must transfer with the part the mark, tag, 
or other record used to comply with this section, unless the part is 
mutilated before it is sold or transferred.

[Doc. No. FAA-2000-8017, 67 FR 2110, Jan. 15, 2002, as amended by Amdt. 
43-38A, 79 FR 67055, Nov. 12, 2014]



Sec. 43.11  Content, form, and disposition of records for inspections
conducted under parts 91 and 125 and Sec. Sec. 135.411(a)(1) and 
135.419 of this chapter.

    (a) Maintenance record entries. The person approving or disapproving 
for return to service an aircraft, airframe, aircraft engine, propeller, 
appliance, or component part after any inspection performed in 
accordance with part 91, 125, Sec. 135.411(a)(1), or Sec. 135.419 
shall make an entry in the maintenance record of that equipment 
containing the following information:
    (1) The type of inspection and a brief description of the extent of 
the inspection.
    (2) The date of the inspection and aircraft total time in service.
    (3) The signature, the certificate number, and kind of certificate 
held by the person approving or disapproving for return to service the 
aircraft, airframe, aircraft engine, propeller, appliance, component 
part, or portions thereof.
    (4) Except for progressive inspections, if the aircraft is found to 
be airworthy and approved for return to service, the following or a 
similarly worded statement--``I certify that this aircraft has been 
inspected in accordance with (insert type) inspection and was determined 
to be in airworthy condition.''
    (5) Except for progressive inspections, if the aircraft is not 
approved for return to service because of needed maintenance, 
noncompliance with applicable specifications, airworthiness directives, 
or other approved data, the following or a similarly worded statement--
``I certify that this aircraft has been inspected in accordance with 
(insert type) inspection and a list of discrepancies and unairworthy 
items dated (date) has been provided for the aircraft owner or 
operator.''
    (6) For progressive inspections, the following or a similarly worded 
statement--``I certify that in accordance with a progressive inspection 
program, a routine inspection of (identify whether aircraft or 
components) and a detailed inspection of (identify components) were 
performed and the (aircraft or components) are (approved or disapproved) 
for return to service.'' If disapproved, the entry will further state 
``and a list of discrepancies and unairworthy items dated (date) has 
been provided to the aircraft owner or operator.''
    (7) If an inspection is conducted under an inspection program 
provided for in part 91, 125, or Sec. 135.411(a)(1), the entry must 
identify the inspection program, that part of the inspection program 
accomplished, and contain a statement that the inspection was performed 
in accordance with the inspections and procedures for that particular 
program.
    (b) Listing of discrepancies and placards. If the person performing 
any inspection required by part 91 or 125 or Sec. 135.411(a)(1) of this 
chapter finds that the aircraft is unairworthy or does not meet the 
applicable type certificate data, airworthiness directives, or other 
approved data upon which its airworthiness depends, that persons must 
give the owner or lessee a signed and dated list of those discrepancies. 
For those items permitted to be inoperative under Sec. 91.213(d)(2) of 
this chapter, that person shall place a placard, that meets the 
aircraft's airworthiness certification regulations, on each inoperative 
instrument and the cockpit control of each item of inoperative 
equipment, marking it ``Inoperative,'' and

[[Page 1021]]

shall add the items to the signed and dated list of discrepancies given 
to the owner or lessee.

[Amdt. 43-23, 47 FR 41085, Sept. 16, 1982, as amended by Amdt. 43-30, 53 
FR 50195, Dec. 13, 1988; Amdt. 43-36, 61 FR 19501, May 1, 1996; 71 FR 
44188, Aug. 4, 2006]



Sec. 43.12  Maintenance records: Falsification, reproduction, 
or alteration.

    (a) No person may make or cause to be made:
    (1) Any fraudulent or intentionally false entry in any record or 
report that is required to be made, kept, or used to show compliance 
with any requirement under this part;
    (2) Any reproduction, for fraudulent purpose, of any record or 
report under this part; or
    (3) Any alteration, for fraudulent purpose, of any record or report 
under this part.
    (b) The commission by any person of an act prohibited under 
paragraph (a) of this section is a basis for suspending or revoking the 
applicable airman, operator, or production certificate, Technical 
Standard Order Authorization, FAA-Parts Manufacturer Approval, or 
Product and Process Specification issued by the Administrator and held 
by that person.

[Amdt. 43-19, 43 FR 22639, May 25, 1978, as amended by Amdt. 43-23, 47 
FR 41085, Sept. 16, 1982]



Sec. 43.13  Performance rules (general).

    (a) Each person performing maintenance, alteration, or preventive 
maintenance on an aircraft, engine, propeller, or appliance shall use 
the methods, techniques, and practices prescribed in the current 
manufacturer's maintenance manual or Instructions for Continued 
Airworthiness prepared by its manufacturer, or other methods, 
techniques, and practices acceptable to the Administrator, except as 
noted in Sec. 43.16. He shall use the tools, equipment, and test 
apparatus necessary to assure completion of the work in accordance with 
accepted industry practices. If special equipment or test apparatus is 
recommended by the manufacturer involved, he must use that equipment or 
apparatus or its equivalent acceptable to the Administrator.
    (b) Each person maintaining or altering, or performing preventive 
maintenance, shall do that work in such a manner and use materials of 
such a quality, that the condition of the aircraft, airframe, aircraft 
engine, propeller, or appliance worked on will be at least equal to its 
original or properly altered condition (with regard to aerodynamic 
function, structural strength, resistance to vibration and 
deterioration, and other qualities affecting airworthiness).
    (c) Special provisions for holders of air carrier operating 
certificates and operating certificates issued under the provisions of 
Part 121 or 135 and Part 129 operators holding operations 
specifications. Unless otherwise notified by the administrator, the 
methods, techniques, and practices contained in the maintenance manual 
or the maintenance part of the manual of the holder of an air carrier 
operating certificate or an operating certificate under Part 121 or 135 
and Part 129 operators holding operations specifications (that is 
required by its operating specifications to provide a continuous 
airworthiness maintenance and inspection program) constitute acceptable 
means of compliance with this section.

[Doc. No. 1993, 29 FR 5451, Apr. 23, 1964, as amended by Amdt. 43-20, 45 
FR 60182, Sept. 11, 1980; Amdt. 43-23, 47 FR 41085, Sept. 16, 1982; 
Amdt. 43-28, 52 FR 20028, June 16, 1987; Amdt. 43-37, 66 FR 21066, Apr. 
27, 2001]



Sec. 43.15  Additional performance rules for inspections.

    (a) General. Each person performing an inspection required by part 
91, 125, or 135 of this chapter, shall--
    (1) Perform the inspection so as to determine whether the aircraft, 
or portion(s) thereof under inspection, meets all applicable 
airworthiness requirements; and
    (2) If the inspection is one provided for in part 125, 135, or Sec. 
91.409(e) of this chapter, perform the inspection in accordance with the 
instructions and procedures set forth in the inspection program for the 
aircraft being inspected.
    (b) Rotorcraft. Each person performing an inspection required by 
Part 91 on a rotorcraft shall inspect the following systems in 
accordance with the maintenance manual or Instructions

[[Page 1022]]

for Continued Airworthiness of the manufacturer concerned:
    (1) The drive shafts or similar systems.
    (2) The main rotor transmission gear box for obvious defects.
    (3) The main rotor and center section (or the equivalent area).
    (4) The auxiliary rotor on helicopters.
    (c) Annual and 100-hour inspections. (1) Each person performing an 
annual or 100-hour inspection shall use a checklist while performing the 
inspection. The checklist may be of the person's own design, one 
provided by the manufacturer of the equipment being inspected or one 
obtained from another source. This checklist must include the scope and 
detail of the items contained in appendix D to this part and paragraph 
(b) of this section.
    (2) Each person approving a reciprocating-engine-powered aircraft 
for return to service after an annual or 100-hour inspection shall, 
before that approval, run the aircraft engine or engines to determine 
satisfactory performance in accordance with the manufacturer's 
recommendations of--
    (i) Power output (static and idle r.p.m.);
    (ii) Magnetos;
    (iii) Fuel and oil pressure; and
    (iv) Cylinder and oil temperature.
    (3) Each person approving a turbine-engine-powered aircraft for 
return to service after an annual, 100-hour, or progressive inspection 
shall, before that approval, run the aircraft engine or engines to 
determine satisfactory performance in accordance with the manufacturer's 
recommendations.
    (d) Progressive inspection. (1) Each person performing a progressive 
inspection shall, at the start of a progressive inspection system, 
inspect the aircraft completely. After this initial inspection, routine 
and detailed inspections must be conducted as prescribed in the 
progressive inspection schedule. Routine inspections consist of visual 
examination or check of the appliances, the aircraft, and its components 
and systems, insofar as practicable without disassembly. Detailed 
inspections consist of a thorough examination of the appliances, the 
aircraft, and its components and systems, with such disassembly as is 
necessary. For the purposes of this subparagraph, the overhaul of a 
component or system is considered to be a detailed inspection.
    (2) If the aircraft is away from the station where inspections are 
normally conducted, an appropriately rated mechanic, a certificated 
repair station, or the manufacturer of the aircraft may perform 
inspections in accordance with the procedures and using the forms of the 
person who would otherwise perform the inspection.

[Doc. No. 1993, 29 FR 5451, Apr. 23, 1964, as amended by Amdt. 43-23, 47 
FR 41086, Sept. 16, 1982; Amdt. 43-25, 51 FR 40702, Nov. 7, 1986; Amdt. 
43-31, 54 FR 34330, Aug. 18, 1989; 71 FR 44188, Aug. 4, 2006]



Sec. 43.16  Airworthiness limitations.

    Each person performing an inspection or other maintenance specified 
in an Airworthiness Limitations section of a manufacturer's maintenance 
manual or Instructions for Continued Airworthiness shall perform the 
inspection or other maintenance in accordance with that section, or in 
accordance with operations specifications approved by the Administrator 
under part 121 or 135, or an inspection program approved under Sec. 
91.409(e).

[71 FR 44188, Aug. 4, 2006]



Sec. 43.17  Maintenance, preventive maintenance, and alterations 
performed on U.S. aeronautical products by certain Canadian persons.

    (a) Definitions. For purposes of this section:
    Aeronautical product means any civil aircraft or airframe, aircraft 
engine, propeller, appliance, component, or part to be installed 
thereon.
    Canadian aeronautical product means any aeronautical product under 
airworthiness regulation by Transport Canada Civil Aviation.
    U.S. aeronautical product means any aeronautical product under 
airworthiness regulation by the FAA.
    (b) Applicability. This section does not apply to any U.S. 
aeronautical products maintained or altered under any bilateral 
agreement made between Canada and any country other than the United 
States.
    (c) Authorized persons. (1) A person holding a valid Transport 
Canada Civil

[[Page 1023]]

Aviation Maintenance Engineer license and appropriate ratings may, with 
respect to a U.S.-registered aircraft located in Canada, perform 
maintenance, preventive maintenance, and alterations in accordance with 
the requirements of paragraph (d) of this section and approve the 
affected aircraft for return to service in accordance with the 
requirements of paragraph (e) of this section.
    (2) A Transport Canada Civil Aviation Approved Maintenance 
Organization (AMO) holding appropriate ratings may, with respect to a 
U.S.-registered aircraft or other U.S. aeronautical products located in 
Canada, perform maintenance, preventive maintenance, and alterations in 
accordance with the requirements of paragraph (d) of this section and 
approve the affected products for return to service in accordance with 
the requirements of paragraph (e) of this section.
    (d) Performance requirements. A person authorized in paragraph (c) 
of this section may perform maintenance (including any inspection 
required by Sec. 91.409 of this chapter, except an annual inspection), 
preventive maintenance, and alterations, provided--
    (1) The person performing the work is authorized by Transport Canada 
Civil Aviation to perform the same type of work with respect to Canadian 
aeronautical products;
    (2) The maintenance, preventive maintenance, or alteration is 
performed in accordance with a Bilateral Aviation Safety Agreement 
between the United States and Canada and associated Maintenance 
Implementation Procedures that provide a level of safety equivalent to 
that provided by the provisions of this chapter;
    (3) The maintenance, preventive maintenance, or alteration is 
performed such that the affected product complies with the applicable 
requirements of part 36 of this chapter; and
    (4) The maintenance, preventive maintenance, or alteration is 
recorded in accordance with a Bilateral Aviation Safety Agreement 
between the United States and Canada and associated Maintenance 
Implementation Procedures that provide a level of safety equivalent to 
that provided by the provisions of this chapter.
    (e) Approval requirements. (1) To return an affected product to 
service, a person authorized in paragraph (c) of this section must 
approve (certify) maintenance, preventive maintenance, and alterations 
performed under this section, except that an Aircraft Maintenance 
Engineer may not approve a major repair or major alteration.
    (2) An AMO whose system of quality control for the maintenance, 
preventive maintenance, alteration, and inspection of aeronautical 
products has been approved by Transport Canada Civil Aviation, or an 
authorized employee performing work for such an AMO, may approve 
(certify) a major repair or major alteration performed under this 
section if the work was performed in accordance with technical data 
approved by the FAA.
    (f) No person may operate in air commerce an aircraft, airframe, 
aircraft engine, propeller, or appliance on which maintenance, 
preventive maintenance, or alteration has been performed under this 
section unless it has been approved for return to service by a person 
authorized in this section.

[Amdt. 43-33, 56 FR 57571, Nov. 12, 1991, as amended by Amdt. 43-40, 71 
FR 40877, July 14, 2005]



   Sec. Appendix A to Part 43--Major Alterations, Major Repairs, and 
                         Preventive Maintenance

    (a) Major alterations--(1) Airframe major alterations. Alterations 
of the following parts and alterations of the following types, when not 
listed in the aircraft specifications issued by the FAA, are airframe 
major alterations:
    (i) Wings.
    (ii) Tail surfaces.
    (iii) Fuselage.
    (iv) Engine mounts.
    (v) Control system.
    (vi) Landing gear.
    (vii) Hull or floats.
    (viii) Elements of an airframe including spars, ribs, fittings, 
shock absorbers, bracing, cowling, fairings, and balance weights.
    (ix) Hydraulic and electrical actuating system of components.
    (x) Rotor blades.
    (xi) Changes to the empty weight or empty balance which result in an 
increase in the maximum certificated weight or center of gravity limits 
of the aircraft.

[[Page 1024]]

    (xii) Changes to the basic design of the fuel, oil, cooling, 
heating, cabin pressurization, electrical, hydraulic, de-icing, or 
exhaust systems.
    (xiii) Changes to the wing or to fixed or movable control surfaces 
which affect flutter and vibration characteristics.
    (2) Powerplant major alterations. The following alterations of a 
powerplant when not listed in the engine specifications issued by the 
FAA, are powerplant major alterations.
    (i) Conversion of an aircraft engine from one approved model to 
another, involving any changes in compression ratio, propeller reduction 
gear, impeller gear ratios or the substitution of major engine parts 
which requires extensive rework and testing of the engine.
    (ii) Changes to the engine by replacing aircraft engine structural 
parts with parts not supplied by the original manufacturer or parts not 
specifically approved by the Administrator.
    (iii) Installation of an accessory which is not approved for the 
engine.
    (iv) Removal of accessories that are listed as required equipment on 
the aircraft or engine specification.
    (v) Installation of structural parts other than the type of parts 
approved for the installation.
    (vi) Conversions of any sort for the purpose of using fuel of a 
rating or grade other than that listed in the engine specifications.
    (3) Propeller major alterations. The following alterations of a 
propeller when not authorized in the propeller specifications issued by 
the FAA are propeller major alterations:
    (i) Changes in blade design.
    (ii) Changes in hub design.
    (iii) Changes in the governor or control design.
    (iv) Installation of a propeller governor or feathering system.
    (v) Installation of propeller de-icing system.
    (vi) Installation of parts not approved for the propeller.
    (4) Appliance major alterations. Alterations of the basic design not 
made in accordance with recommendations of the appliance manufacturer or 
in accordance with an FAA Airworthiness Directive are appliance major 
alterations. In addition, changes in the basic design of radio 
communication and navigation equipment approved under type certification 
or a Technical Standard Order that have an effect on frequency 
stability, noise level, sensitivity, selectivity, distortion, spurious 
radiation, AVC characteristics, or ability to meet environmental test 
conditions and other changes that have an effect on the performance of 
the equipment are also major alterations.
    (b) Major repairs--(1) Airframe major repairs. Repairs to the 
following parts of an airframe and repairs of the following types, 
involving the strengthening, reinforcing, splicing, and manufacturing of 
primary structural members or their replacement, when replacement is by 
fabrication such as riveting or welding, are airframe major repairs.
    (i) Box beams.
    (ii) Monocoque or semimonocoque wings or control surfaces.
    (iii) Wing stringers or chord members.
    (iv) Spars.
    (v) Spar flanges.
    (vi) Members of truss-type beams.
    (vii) Thin sheet webs of beams.
    (viii) Keel and chine members of boat hulls or floats.
    (ix) Corrugated sheet compression members which act as flange 
material of wings or tail surfaces.
    (x) Wing main ribs and compression members.
    (xi) Wing or tail surface brace struts.
    (xii) Engine mounts.
    (xiii) Fuselage longerons.
    (xiv) Members of the side truss, horizontal truss, or bulkheads.
    (xv) Main seat support braces and brackets.
    (xvi) Landing gear brace struts.
    (xvii) Axles.
    (xviii) Wheels.
    (xix) Skis, and ski pedestals.
    (xx) Parts of the control system such as control columns, pedals, 
shafts, brackets, or horns.
    (xxi) Repairs involving the substitution of material.
    (xxii) The repair of damaged areas in metal or plywood stressed 
covering exceeding six inches in any direction.
    (xxiii) The repair of portions of skin sheets by making additional 
seams.
    (xxiv) The splicing of skin sheets.
    (xxv) The repair of three or more adjacent wing or control surface 
ribs or the leading edge of wings and control surfaces, between such 
adjacent ribs.
    (xxvi) Repair of fabric covering involving an area greater than that 
required to repair two adjacent ribs.
    (xxvii) Replacement of fabric on fabric covered parts such as wings, 
fuselages, stabilizers, and control surfaces.
    (xxviii) Repairing, including rebottoming, of removable or integral 
fuel tanks and oil tanks.
    (2) Powerplant major repairs. Repairs of the following parts of an 
engine and repairs of the following types, are powerplant major repairs:
    (i) Separation or disassembly of a crankcase or crankshaft of a 
reciprocating engine equipped with an integral supercharger.
    (ii) Separation or disassembly of a crankcase or crankshaft of a 
reciprocating engine equipped with other than spur-type propeller 
reduction gearing.

[[Page 1025]]

    (iii) Special repairs to structural engine parts by welding, 
plating, metalizing, or other methods.
    (3) Propeller major repairs. Repairs of the following types to a 
propeller are propeller major repairs:
    (i) Any repairs to, or straightening of steel blades.
    (ii) Repairing or machining of steel hubs.
    (iii) Shortening of blades.
    (iv) Retipping of wood propellers.
    (v) Replacement of outer laminations on fixed pitch wood propellers.
    (vi) Repairing elongated bolt holes in the hub of fixed pitch wood 
propellers.
    (vii) Inlay work on wood blades.
    (viii) Repairs to composition blades.
    (ix) Replacement of tip fabric.
    (x) Replacement of plastic covering.
    (xi) Repair of propeller governors.
    (xii) Overhaul of controllable pitch propellers.
    (xiii) Repairs to deep dents, cuts, scars, nicks, etc., and 
straightening of aluminum blades.
    (xiv) The repair or replacement of internal elements of blades.
    (4) Appliance major repairs. Repairs of the following types to 
appliances are appliance major repairs:
    (i) Calibration and repair of instruments.
    (ii) Calibration of radio equipment.
    (iii) Rewinding the field coil of an electrical accessory.
    (iv) Complete disassembly of complex hydraulic power valves.
    (v) Overhaul of pressure type carburetors, and pressure type fuel, 
oil and hydraulic pumps.
    (c) Preventive maintenance. Preventive maintenance is limited to the 
following work, provided it does not involve complex assembly 
operations:
    (1) Removal, installation, and repair of landing gear tires.
    (2) Replacing elastic shock absorber cords on landing gear.
    (3) Servicing landing gear shock struts by adding oil, air, or both.
    (4) Servicing landing gear wheel bearings, such as cleaning and 
greasing.
    (5) Replacing defective safety wiring or cotter keys.
    (6) Lubrication not requiring disassembly other than removal of 
nonstructural items such as cover plates, cowlings, and fairings.
    (7) Making simple fabric patches not requiring rib stitching or the 
removal of structural parts or control surfaces. In the case of 
balloons, the making of small fabric repairs to envelopes (as defined 
in, and in accordance with, the balloon manufacturers' instructions) not 
requiring load tape repair or replacement.
    (8) Replenishing hydraulic fluid in the hydraulic reservoir.
    (9) Refinishing decorative coating of fuselage, balloon baskets, 
wings tail group surfaces (excluding balanced control surfaces), 
fairings, cowlings, landing gear, cabin, or cockpit interior when 
removal or disassembly of any primary structure or operating system is 
not required.
    (10) Applying preservative or protective material to components 
where no disassembly of any primary structure or operating system is 
involved and where such coating is not prohibited or is not contrary to 
good practices.
    (11) Repairing upholstery and decorative furnishings of the cabin, 
cockpit, or balloon basket interior when the repairing does not require 
disassembly of any primary structure or operating system or interfere 
with an operating system or affect the primary structure of the 
aircraft.
    (12) Making small simple repairs to fairings, nonstructural cover 
plates, cowlings, and small patches and reinforcements not changing the 
contour so as to interfere with proper air flow.
    (13) Replacing side windows where that work does not interfere with 
the structure or any operating system such as controls, electrical 
equipment, etc.
    (14) Replacing safety belts.
    (15) Replacing seats or seat parts with replacement parts approved 
for the aircraft, not involving disassembly of any primary structure or 
operating system.
    (16) Trouble shooting and repairing broken circuits in landing light 
wiring circuits.
    (17) Replacing bulbs, reflectors, and lenses of position and landing 
lights.
    (18) Replacing wheels and skis where no weight and balance 
computation is involved.
    (19) Replacing any cowling not requiring removal of the propeller or 
disconnection of flight controls.
    (20) Replacing or cleaning spark plugs and setting of spark plug gap 
clearance.
    (21) Replacing any hose connection except hydraulic connections.
    (22) Replacing prefabricated fuel lines.
    (23) Cleaning or replacing fuel and oil strainers or filter 
elements.
    (24) Replacing and servicing batteries.
    (25) Cleaning of balloon burner pilot and main nozzles in accordance 
with the balloon manufacturer's instructions.
    (26) Replacement or adjustment of nonstructural standard fasteners 
incidental to operations.
    (27) The interchange of balloon baskets and burners on envelopes 
when the basket or burner is designated as interchangeable in the 
balloon type certificate data and the baskets and burners are 
specifically designed for quick removal and installation.
    (28) The installations of anti-misfueling devices to reduce the 
diameter of fuel tank filler openings provided the specific device has 
been made a part of the aircraft type

[[Page 1026]]

certificiate data by the aircraft manufacturer, the aircraft 
manufacturer has provided FAA-approved instructions for installation of 
the specific device, and installation does not involve the disassembly 
of the existing tank filler opening.
    (29) Removing, checking, and replacing magnetic chip detectors.
    (30) The inspection and maintenance tasks prescribed and 
specifically identified as preventive maintenance in a primary category 
aircraft type certificate or supplemental type certificate holder's 
approved special inspection and preventive maintenance program when 
accomplished on a primary category aircraft provided:
    (i) They are performed by the holder of at least a private pilot 
certificate issued under part 61 who is the registered owner (including 
co-owners) of the affected aircraft and who holds a certificate of 
competency for the affected aircraft (1) issued by a school approved 
under Sec. 147.21(e) of this chapter; (2) issued by the holder of the 
production certificate for that primary category aircraft that has a 
special training program approved under Sec. 21.24 of this subchapter; 
or (3) issued by another entity that has a course approved by the 
Administrator; and
    (ii) The inspections and maintenance tasks are performed in 
accordance with instructions contained by the special inspection and 
preventive maintenance program approved as part of the aircraft's type 
design or supplemental type design.
    (31) Removing and replacing self-contained, front instrument panel-
mounted navigation and communication devices that employ tray-mounted 
connectors that connect the unit when the unit is installed into the 
instrument panel, (excluding automatic flight control systems, 
transponders, and microwave frequency distance measuring equipment 
(DME)). The approved unit must be designed to be readily and repeatedly 
removed and replaced, and pertinent instructions must be provided. Prior 
to the unit's intended use, and operational check must be performed in 
accordance with the applicable sections of part 91 of this chapter.

(Secs. 313, 601 through 610, and 1102, Federal Aviation Act of 1958 as 
amended (49 U.S.C. 1354, 1421 through 1430 and 1502); (49 U.S.C. 106(g) 
(Revised Pub. L. 97-449, Jan. 21, 1983); and 14 CFR 11.45)

[Doc. No. 1993, 29 FR 5451, Apr. 23, 1964, as amended by Amdt. 43-14, 37 
FR 14291, June 19, 1972; Amdt. 43-23, 47 FR 41086, Sept. 16, 1982; Amdt. 
43-24, 49 FR 44602, Nov. 7, 1984; Amdt. 43-25, 51 FR 40703, Nov. 7, 
1986; Amdt. 43-27, 52 FR 17277, May 6, 1987; Amdt. 43-34, 57 FR 41369, 
Sept. 9, 1992; Amdt. 43-36, 61 FR 19501, May 1, 1996; Amdt. 43-45, 77 FR 
71096, Nov. 29, 2012]



    Sec. Appendix B to Part 43--Recording of Major Repairs and Major 
                               Alterations

    (a) Except as provided in paragraphs (b), (c), and (d) of this 
appendix, each person performing a major repair or major alteration 
shall--
    (1) Execute FAA Form 337 at least in duplicate;
    (2) Give a signed copy of that form to the aircraft owner; and
    (3) Forward a copy of that form to the FAA Aircraft Registration 
Branch in Oklahoma City, Oklahoma, within 48 hours after the aircraft, 
airframe, aircraft engine, propeller, or appliance is approved for 
return to service.
    (b) For major repairs made in accordance with a manual or 
specifications acceptable to the Administrator, a certificated repair 
station may, in place of the requirements of paragraph (a)--
    (1) Use the customer's work order upon which the repair is recorded;
    (2) Give the aircraft owner a signed copy of the work order and 
retain a duplicate copy for at least two years from the date of approval 
for return to service of the aircraft, airframe, aircraft engine, 
propeller, or appliance;
    (3) Give the aircraft owner a maintenance release signed by an 
authorized representative of the repair station and incorporating the 
following information:
    (i) Identity of the aircraft, airframe, aircraft engine, propeller 
or appliance.
    (ii) If an aircraft, the make, model, serial number, nationality and 
registration marks, and location of the repaired area.
    (iii) If an airframe, aircraft engine, propeller, or appliance, give 
the manufacturer's name, name of the part, model, and serial numbers (if 
any); and
    (4) Include the following or a similarly worded statement--
    ``The aircraft, airframe, aircraft engine, propeller, or appliance 
identified above was repaired and inspected in accordance with current 
Regulations of the Federal Aviation Agency and is approved for return to 
service.
    Pertinent details of the repair are on file at this repair station 
under Order No. ------,

Date____________________________________________________________________
Signed__________________________________________________________________

For signature of authorized representative)

Repair station name) (Certificate No.)
------------------------.''

(Address)
    (c) Except as provided in paragraph (d) of this appendix, for a 
major repair or major alteration made by a person authorized in Sec. 
43.17, the person who performs the major repair or major alteration and 
the person authorized by Sec. 43.17 to approve that work shall

[[Page 1027]]

execute an FAA Form 337 at least in duplicate. A completed copy of that 
form shall be--
    (1) Given to the aircraft owner; and
    (2) Forwarded to the Federal Aviation Administration, Aircraft 
Registration Branch, AFS-750, Post Office Box 25504, Oklahoma City, OK 
73125, within 48 hours after the work is inspected.
    (d) For extended-range fuel tanks installed within the passenger 
compartment or a baggage compartment, the person who performs the work 
and the person authorized to approve the work by Sec. 43.7 shall 
execute an FAA Form 337 in at least triplicate. A completed copy of that 
form shall be--
    (1) Placed on board the aircraft as specified in Sec. 91.417 of 
this chapter;
    (2) Given to the aircraft owner; and
    (3) Forwarded to the Federal Aviation Administration, Aircraft 
Registration Branch, AFS-751, Post Office Box 25724, Oklahoma City, OK 
73125, within 48 hours after the work is inspected.

(Secs. 101, 610, 72 Stat. 737, 780, 49 U.S.C. 1301, 1430)

[Doc. No. 1993, 29 FR 5451, Apr. 23, 1964, as amended by Amdt. 43-10, 33 
FR 15989, Oct. 31, 1968; Amdt. 43-29, 52 FR 34101, Sept. 9, 1987; Amdt. 
43-31, 54 FR 34330, Aug. 18, 1989; 71 FR 58495, Oct. 4, 2006; Amdt. 43-
41, 72 FR 53680, Sept. 20, 2007]



                  Sec. Appendix C to Part 43 [Reserved]



Sec. Appendix D to Part 43--Scope and Detail of Items (as Applicable to 
     the Particular Aircraft) To Be Included in Annual and 100-Hour 
                               Inspections

    (a) Each person performing an annual or 100-hour inspection shall, 
before that inspection, remove or open all necessary inspection plates, 
access doors, fairing, and cowling. He shall thoroughly clean the 
aircraft and aircraft engine.
    (b) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) the following components of the fuselage and 
hull group:
    (1) Fabric and skin--for deterioration, distortion, other evidence 
of failure, and defective or insecure attachment of fittings.
    (2) Systems and components--for improper installation, apparent 
defects, and unsatisfactory operation.
    (3) Envelope, gas bags, ballast tanks, and related parts--for poor 
condition.
    (c) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) the following components of the cabin and 
cockpit group:
    (1) Generally--for uncleanliness and loose equipment that might foul 
the controls.
    (2) Seats and safety belts--for poor condition and apparent defects.
    (3) Windows and windshields--for deterioration and breakage.
    (4) Instruments--for poor condition, mounting, marking, and (where 
practicable) improper operation.
    (5) Flight and engine controls--for improper installation and 
improper operation.
    (6) Batteries--for improper installation and improper charge.
    (7) All systems--for improper installation, poor general condition, 
apparent and obvious defects, and insecurity of attachment.
    (d) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) components of the engine and nacelle group as 
follows:
    (1) Engine section--for visual evidence of excessive oil, fuel, or 
hydraulic leaks, and sources of such leaks.
    (2) Studs and nuts--for improper torquing and obvious defects.
    (3) Internal engine--for cylinder compression and for metal 
particles or foreign matter on screens and sump drain plugs. If there is 
weak cylinder compression, for improper internal condition and improper 
internal tolerances.
    (4) Engine mount--for cracks, looseness of mounting, and looseness 
of engine to mount.
    (5) Flexible vibration dampeners--for poor condition and 
deterioration.
    (6) Engine controls--for defects, improper travel, and improper 
safetying.
    (7) Lines, hoses, and clamps--for leaks, improper condition and 
looseness.
    (8) Exhaust stacks--for cracks, defects, and improper attachment.
    (9) Accessories--for apparent defects in security of mounting.
    (10) All systems--for improper installation, poor general condition, 
defects, and insecure attachment.
    (11) Cowling--for cracks, and defects.
    (e) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) the following components of the landing gear 
group:
    (1) All units--for poor condition and insecurity of attachment.
    (2) Shock absorbing devices--for improper oleo fluid level.
    (3) Linkages, trusses, and members--for undue or excessive wear 
fatigue, and distortion.
    (4) Retracting and locking mechanism--for improper operation.
    (5) Hydraulic lines--for leakage.
    (6) Electrical system--for chafing and improper operation of 
switches.
    (7) Wheels--for cracks, defects, and condition of bearings.
    (8) Tires--for wear and cuts.
    (9) Brakes--for improper adjustment.
    (10) Floats and skis--for insecure attachment and obvious or 
apparent defects.

[[Page 1028]]

    (f) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) all components of the wing and center section 
assembly for poor general condition, fabric or skin deterioration, 
distortion, evidence of failure, and insecurity of attachment.
    (g) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) all components and systems that make up the 
complete empennage assembly for poor general condition, fabric or skin 
deterioration, distortion, evidence of failure, insecure attachment, 
improper component installation, and improper component operation.
    (h) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) the following components of the propeller 
group:
    (1) Propeller assembly--for cracks, nicks, binds, and oil leakage.
    (2) Bolts--for improper torquing and lack of safetying.
    (3) Anti-icing devices--for improper operations and obvious defects.
    (4) Control mechanisms--for improper operation, insecure mounting, 
and restricted travel.
    (i) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) the following components of the radio group:
    (1) Radio and electronic equipment--for improper installation and 
insecure mounting.
    (2) Wiring and conduits--for improper routing, insecure mounting, 
and obvious defects.
    (3) Bonding and shielding--for improper installation and poor 
condition.
    (4) Antenna including trailing antenna--for poor condition, insecure 
mounting, and improper operation.
    (j) Each person performing an annual or 100-hour inspection shall 
inspect (where applicable) each installed miscellaneous item that is not 
otherwise covered by this listing for improper installation and improper 
operation.



    Sec. Appendix E to Part 43--Altimeter System Test and Inspection

    Each person performing the altimeter system tests and inspections 
required by Sec. 91.411 shall comply with the following:
    (a) Static pressure system:
    (1) Ensure freedom from entrapped moisture and restrictions.
    (2) Determine that leakage is within the tolerances established in 
Sec. 23.1325 or Sec. 25.1325, whichever is applicable.
    (3) Determine that the static port heater, if installed, is 
operative.
    (4) Ensure that no alterations or deformations of the airframe 
surface have been made that would affect the relationship between air 
pressure in the static pressure system and true ambient static air 
pressure for any flight condition.
    (b) Altimeter:
    (1) Test by an appropriately rated repair facility in accordance 
with the following subparagraphs. Unless otherwise specified, each test 
for performance may be conducted with the instrument subjected to 
vibration. When tests are conducted with the temperature substantially 
different from ambient temperature of approximately 25 degrees C., 
allowance shall be made for the variation from the specified condition.
    (i) Scale error. With the barometric pressure scale at 29.92 inches 
of mercury, the altimeter shall be subjected successively to pressures 
corresponding to the altitude specified in Table I up to the maximum 
normally expected operating altitude of the airplane in which the 
altimeter is to be installed. The reduction in pressure shall be made at 
a rate not in excess of 20,000 feet per minute to within approximately 
2,000 feet of the test point. The test point shall be approached at a 
rate compatible with the test equipment. The altimeter shall be kept at 
the pressure corresponding to each test point for at least 1 minute, but 
not more than 10 minutes, before a reading is taken. The error at all 
test points must not exceed the tolerances specified in Table I.
    (ii) Hysteresis. The hysteresis test shall begin not more than 15 
minutes after the altimeter's initial exposure to the pressure 
corresponding to the upper limit of the scale error test prescribed in 
subparagraph (i); and while the altimeter is at this pressure, the 
hysteresis test shall commence. Pressure shall be increased at a rate 
simulating a descent in altitude at the rate of 5,000 to 20,000 feet per 
minute until within 3,000 feet of the first test point (50 percent of 
maximum altitude). The test point shall then be approached at a rate of 
approximately 3,000 feet per minute. The altimeter shall be kept at this 
pressure for at least 5 minutes, but not more than 15 minutes, before 
the test reading is taken. After the reading has been taken, the 
pressure shall be increased further, in the same manner as before, until 
the pressure corresponding to the second test point (40 percent of 
maximum altitude) is reached. The altimeter shall be kept at this 
pressure for at least 1 minute, but not more than 10 minutes, before the 
test reading is taken. After the reading has been taken, the pressure 
shall be increased further, in the same manner as before, until 
atmospheric pressure is reached. The reading of the altimeter at either 
of the two test points shall not differ by more than the tolerance 
specified in Table II from the reading of the altimeter for the 
corresponding altitude recorded during the scale error test prescribed 
in paragraph (b)(i).

[[Page 1029]]

    (iii) After effect. Not more than 5 minutes after the completion of 
the hysteresis test prescribed in paragraph (b)(ii), the reading of the 
altimeter (corrected for any change in atmospheric pressure) shall not 
differ from the original atmospheric pressure reading by more than the 
tolerance specified in Table II.
    (iv) Friction. The altimeter shall be subjected to a steady rate of 
decrease of pressure approximating 750 feet per minute. At each altitude 
listed in Table III, the change in reading of the pointers after 
vibration shall not exceed the corresponding tolerance listed in Table 
III.
    (v) Case leak. The leakage of the altimeter case, when the pressure 
within it corresponds to an altitude of 18,000 feet, shall not change 
the altimeter reading by more than the tolerance shown in Table II 
during an interval of 1 minute.
    (vi) Barometric scale error. At constant atmospheric pressure, the 
barometric pressure scale shall be set at each of the pressures (falling 
within its range of adjustment) that are listed in Table IV, and shall 
cause the pointer to indicate the equivalent altitude difference shown 
in Table IV with a tolerance of 25 feet.
    (2) Altimeters which are the air data computer type with associated 
computing systems, or which incorporate air data correction internally, 
may be tested in a manner and to specifications developed by the 
manufacturer which are acceptable to the Administrator.
    (c) Automatic Pressure Altitude Reporting Equipment and ATC 
Transponder System Integration Test. The test must be conducted by an 
appropriately rated person under the conditions specified in paragraph 
(a). Measure the automatic pressure altitude at the output of the 
installed ATC transponder when interrogated on Mode C at a sufficient 
number of test points to ensure that the altitude reporting equipment, 
altimeters, and ATC transponders perform their intended functions as 
installed in the aircraft. The difference between the automatic 
reporting output and the altitude displayed at the altimeter shall not 
exceed 125 feet.
    (d) Records: Comply with the provisions of Sec. 43.9 of this 
chapter as to content, form, and disposition of the records. The person 
performing the altimeter tests shall record on the altimeter the date 
and maximum altitude to which the altimeter has been tested and the 
persons approving the airplane for return to service shall enter that 
data in the airplane log or other permanent record.

                                 Table I
------------------------------------------------------------------------
                                      Equivalent
                                       pressure      Tolerance (feet)
                                       mercury)
------------------------------------------------------------------------
-1,000..............................      31.018                20
0...................................      29.921                20
500.................................      29.385                20
1,000...............................      28.856                20
1,500...............................      28.335                25
2,000...............................      27.821                30
3,000...............................      26.817                30
4,000...............................      25.842                35
6,000...............................      23.978                40
8,000...............................      22.225                60
10,000..............................      20.577                80
12,000..............................      19.029                90
14,000..............................      17.577               100
16,000..............................      16.216               110
18,000..............................      14.942               120
20,000..............................      13.750               130
22,000..............................      12.636               140
25,000..............................      11.104               155
30,000..............................       8.885               180
35,000..............................       7.041               205
40,000..............................       5.538               230
45,000..............................       4.355               255
50,000..............................       3.425               280
------------------------------------------------------------------------


                        Table II--Test Tolerances
------------------------------------------------------------------------
                                                              Tolerance
                           Test                                (feet)
------------------------------------------------------------------------
Case Leak Test............................................        
    Sec. Appendix F to Part 43--ATC Transponder Tests and Inspections

    The ATC transponder tests required by Sec. 91.413 of this chapter 
may be conducted using a bench check or portable test equipment and must 
meet the requirements prescribed in paragraphs (a) through (j) of this 
appendix. If portable test equipment with appropriate coupling to the 
aircraft antenna system is used, operate the test equipment for ATCRBS 
transponders at a nominal rate of 235 interrogations per second to avoid 
possible ATCRBS interference. Operate the test equipment at a nominal 
rate of 50 Mode S interrogations per second for Mode S. An additional 3 
dB loss is allowed to compensate for antenna coupling errors during 
receiver sensitivity measurements conducted in accordance with paragraph 
(c)(1) when using portable test equipment.
    (a) Radio Reply Frequency:
    (1) For all classes of ATCRBS transponders, interrogate the 
transponder and verify that the reply frequency is 1090 3 Megahertz (MHz).
    (2) For classes 1B, 2B, and 3B Mode S transponders, interrogate the 
transponder and verify that the reply frequency is 1090 3 MHz.
    (3) For classes 1B, 2B, and 3B Mode S transponders that incorporate 
the optional 1090 1 MHz reply frequency, 
interrogate the transponder and verify that the reply frequency is 
correct.
    (4) For classes 1A, 2A, 3A, and 4 Mode S transponders, interrogate 
the transponder and verify that the reply frequency is 1090 1 MHz.
    (b) Suppression: When Classes 1B and 2B ATCRBS Transponders, or 
Classes 1B, 2B, and 3B Mode S transponders are interrogated Mode 3/A at 
an interrogation rate between 230 and 1,000 interrogations per second; 
or when Classes 1A and 2A ATCRBS Transponders, or Classes 1B, 2A, 3A, 
and 4 Mode S transponders are interrogated at a rate between 230 and 
1,200 Mode 3/A interrogations per second:
    (1) Verify that the transponder does not respond to more than 1 
percent of ATCRBS interrogations when the amplitude of P2 
pulse is equal to the P1 pulse.
    (2) Verify that the transponder replies to at least 90 percent of 
ATCRBS interrogations when the amplitude of the P2 pulse is 9 
dB less than the P1 pulse. If the test is conducted with a 
radiated test signal, the interrogation rate shall be 235 5 interrogations per second unless a higher rate has 
been approved for the test equipment used at that location.
    (c) Receiver Sensitivity:
    (1) Verify that for any class of ATCRBS Transponder, the receiver 
minimum triggering level (MTL) of the system is -73 4 dbm, or that for any class of Mode S transponder the 
receiver MTL for Mode S format (P6 type) interrogations is -74 3 dbm by use of a test set either:
    (i) Connected to the antenna end of the transmission line;
    (ii) Connected to the antenna terminal of the transponder with a 
correction for transmission line loss; or
    (iii) Utilized radiated signal.
    (2) Verify that the difference in Mode 3/A and Mode C receiver 
sensitivity does not exceed 1 db for either any class of ATCRBS 
transponder or any class of Mode S transponder.
    (d) Radio Frequency (RF) Peak Output Power:
    (1) Verify that the transponder RF output power is within 
specifications for the class of transponder. Use the same conditions as 
described in (c)(1)(i), (ii), and (iii) above.
    (i) For Class 1A and 2A ATCRBS transponders, verify that the minimum 
RF peak output power is at least 21.0 dbw (125 watts).
    (ii) For Class 1B and 2B ATCRBS Transponders, verify that the 
minimum RF peak output power is at least 18.5 dbw (70 watts).
    (iii) For Class 1A, 2A, 3A, and 4 and those Class 1B, 2B, and 3B 
Mode S transponders that include the optional high RF peak output power, 
verify that the minimum RF peak output power is at least 21.0 dbw (125 
watts).
    (iv) For Classes 1B, 2B, and 3B Mode S transponders, verify that the 
minimum RF peak output power is at least 18.5 dbw (70 watts).
    (v) For any class of ATCRBS or any class of Mode S transponders, 
verify that the maximum RF peak output power does not exceed 27.0 dbw 
(500 watts).

    Note: The tests in (e) through (j) apply only to Mode S 
transponders.

    (e) Mode S Diversity Transmission Channel Isolation: For any class 
of Mode S transponder that incorporates diversity operation, verify that 
the RF peak output power transmitted from the selected antenna exceeds

[[Page 1031]]

the power transmitted from the nonselected antenna by at least 20 db.
    (f) Mode S Address: Interrogate the Mode S transponder and verify 
that it replies only to its assigned address. Use the correct address 
and at least two incorrect addresses. The interrogations should be made 
at a nominal rate of 50 interrogations per second.
    (g) Mode S Formats: Interrogate the Mode S transponder with uplink 
formats (UF) for which it is equipped and verify that the replies are 
made in the correct format. Use the surveillance formats UF=4 and 5. 
Verify that the altitude reported in the replies to UF=4 are the same as 
that reported in a valid ATCRBS Mode C reply. Verify that the identity 
reported in the replies to UF=5 are the same as that reported in a valid 
ATCRBS Mode 3/A reply. If the transponder is so equipped, use the 
communication formats UF=20, 21, and 24.
    (h) Mode S All-Call Interrogations: Interrogate the Mode S 
transponder with the Mode S-only all-call format UF=11, and the ATCRBS/
Mode S all-call formats (1.6 microsecond P4 pulse) and verify 
that the correct address and capability are reported in the replies 
(downlink format DF=11).
    (i) ATCRBS-Only All-Call Interrogation: Interrogate the Mode S 
transponder with the ATCRBS-only all-call interrogation (0.8 microsecond 
P4 pulse) and verify that no reply is generated.
    (j) Squitter: Verify that the Mode S transponder generates a correct 
squitter approximately once per second.
    (k) Records: Comply with the provisions of Sec. 43.9 of this 
chapter as to content, form, and disposition of the records.

[Amdt. 43-26, 52 FR 3390, Feb. 3, 1987; 52 FR 6651, Mar. 4, 1987, as 
amended by Amdt. 43-31, 54 FR 34330, Aug. 18, 1989]



PART 45_IDENTIFICATION AND REGISTRATION MARKING--Table of Contents



                            Subpart A_General

Sec.
45.1 Applicability.

               Subpart B_Marking of Products and Articles

45.10 Marking.
45.11 Marking of products.
45.13 Identification data.
45.15 Marking requirements for PMA articles, TSO articles, and Critical 
          parts.
45.16 Marking of life-limited parts.

              Subpart C_Nationality and Registration Marks

45.21 General.
45.22 Exhibition, antique, and other aircraft: Special rules.
45.23 Display of marks; general.
45.25 Location of marks on fixed-wing aircraft.
45.27 Location of marks; nonfixed-wing aircraft.
45.29 Size of marks.
45.31 Marking of export aircraft.
45.33 Sale of aircraft; removal of marks.

    Authority: 49 U.S.C. 106(g), 40103, 40113-40114, 44101-44105, 44107-
44111, 44504, 44701, 44708-44709, 44711-44713, 44725, 45302-45303, 
46104, 46304, 46306, 47122.

    Source: Docket No. 2047, 29 FR 3223, Mar. 11, 1964, unless otherwise 
noted.

    Editorial Note: Nomenclature changes to part 45 appear at 74 FR 
53394, Oct. 16, 2009.



                            Subpart A_General



Sec. 45.1  Applicability.

    This part prescribes the requirements for--
    (a) Marking products and articles manufactured under--
    (1) A type certificate;
    (2) A production approval as defined under part 21 of this chapter; 
and
    (3) The provisions of an agreement between the United States and 
another country or jurisdiction for the acceptance of products and 
articles; and
    (b) Nationality and registration marking of U.S. registered 
aircraft.

[Doc. No. 2047, 29 FR 3223, Mar. 11, 1964, as amended by Amdt. 45-3, 32 
FR 188, Jan. 10, 1967; Amdt. 45-26, 74 FR 53394, Oct. 16, 2009]



               Subpart B_Marking of Products and Articles



Sec. 45.10  Marking.

    No person may mark a product or article in accordance with this 
subpart unless--
    (a) That person produced the product or article --
    (1) Under part 21, subpart F, G, K, or O of this chapter; or
    (2) For export to the United States under the provisions of an 
agreement between the United States and another country or jurisdiction 
for the acceptance of products and articles; and
    (b) That product or article conforms to its approved design, and is 
in a condition for safe operation; and, for a

[[Page 1032]]

TSO article; that TSO article meets the applicable performance 
standards.

[Doc. No. FAA-2006-25877, Amdt. 45-26, 74 FR 53394, Oct. 16, 2009]



Sec. 45.11  Marking of products.

    (a) Aircraft. A manufacturer of aircraft covered under Sec. 21.182 
of this chapter must mark each aircraft by attaching a fireproof 
identification plate that--
    (1) Includes the information specified in Sec. 45.13 using an 
approved method of fireproof marking;
    (2) Must be secured in such a manner that it will not likely be 
defaced or removed during normal service, or lost or destroyed in an 
accident; and
    (3) Except as provided in paragraphs (d) through (h) of this 
section, must be secured to the aircraft fuselage exterior so that it is 
legible to a person on the ground, and must be either adjacent to and 
aft of the rear-most entrance door or on the fuselage surface near the 
tail surfaces.
    (b) Aircraft engines. A manufacturer of an aircraft engine produced 
under a type certificate or production certificate must mark each engine 
by attaching a fireproof identification plate. Such plate--
    (1) Must include the information specified in Sec. 45.13 using an 
approved method of fireproof marking;
    (2) Must be affixed to the engine at an accessible location; and
    (3) Must be secured in such a manner that it will not likely be 
defaced or removed during normal service, or lost or destroyed in an 
accident.
    (c) Propellers and propeller blades and hubs. Each person who 
produces a propeller, propeller blade, or propeller hub under a type 
certificate or production certificate must mark each product or part 
using an approved fireproof method. The marking must--
    (1) Be placed on a non-critical surface;
    (2) Contain the information specified in Sec. 45.13;
    (3) Not likely be defaced or removed during normal service; and
    (4) Not likely be lost or destroyed in an accident.
    (d) Manned free balloons. A manufacturer of manned free balloons 
must mark each balloon by attaching the identification plate described 
in paragraph (a) of this section. The plate must be secured to the 
balloon envelope and must be located, if practicable, where it is 
legible to the operator when the balloon is inflated. In addition, the 
basket and heater assembly must be permanently and legibly marked with 
the manufacturer's name, part number (or equivalent), and serial number 
(or equivalent).
    (e) Aircraft manufactured before March 7, 1988. The owner or 
operator of an aircraft manufactured before March 7, 1988 must mark the 
aircraft by attaching the identification plate required by paragraph (a) 
of this section. The plate must be secured at an accessible exterior or 
interior location near an entrance, if the model designation and 
builder's serial number are also displayed on the exterior of the 
aircraft fuselage. The model designation and builder's serial number 
must be--
    (1) Legible to a person on the ground,
    (2) Located either adjacent to and aft of the rear-most entrance 
door or on the fuselage near the tail surfaces, and
    (3) Displayed in such a manner that they are not likely to be 
defaced or removed during normal service.
    (f) For powered parachutes and weight-shift-control aircraft, the 
identification plate required by paragraph (a) of this section must be 
secured to the exterior of the aircraft fuselage so that it is legible 
to a person on the ground.
    (g) The identification plate described in paragraph (a) of this 
section may be secured to the aircraft at an accessible location near an 
entrance for--
    (1) Aircraft produced for--
    (i) Operations under part 121 of this chapter,
    (ii) Commuter operations (as defined in Sec. 110.2 of this 
chapter), or
    (iii) Export.
    (2) Aircraft operating under part 121 of this chapter and under an 
FAA-approved continuous airworthiness maintenance program; or
    (3) Aircraft operating in commuter air carrier operations (as 
defined in Sec. 110.2 of this chapter) under an FAA-approved continuous 
airworthiness maintenance program.

[[Page 1033]]

    (h) Gliders. Paragraphs (a)(3) and (e) of this section do not apply 
to gliders.

[Amdt. 45-26, 74 FR 53394, Oct. 16, 2009, as amended by Amdt. 45-27, 76 
FR 7486, Feb. 10, 2011]



Sec. 45.13  Identification data.

    (a) The identification required by Sec. 45.11 (a) through (c) must 
include the following information:
    (1) Builder's name.
    (2) Model designation.
    (3) Builder's serial number.
    (4) Type certificate number, if any.
    (5) Production certificate number, if any.
    (6) For aircraft engines, the established rating.
    (7) On or after January 1, 1984, for aircraft engines specified in 
part 34 of this chapter, the date of manufacture as defined in Sec. 
34.1 of this chapter, and a designation, approved by the FAA, that 
indicates compliance with the applicable exhaust emission provisions of 
part 34 of this chapter and 40 CFR part 87. Approved designations 
include COMPLY, EXEMPT, and NON-US, as appropriate. After December 31, 
2012, approved designations also include EXEMPT NEW, and EXCEPTED SPARE, 
as appropriate.
    (i) The designation COMPLY indicates that the engine is in 
compliance with all of the applicable exhaust emissions provisions of 
part 34. For any engine with a rated thrust in excess of 26.7 
kilonewtons (6000 pounds) which is not used or intended for use in 
commercial operations and which is in compliance with the applicable 
provisions of part 34, but does not comply with the hydrocarbon 
emissions standard of Sec. 34.21(d), the statement ``May not be used as 
a commercial aircraft engine'' must be noted in the permanent powerplant 
record that accompanies the engine at the time of manufacture of the 
engine.
    (ii) The designation EXEMPT indicates that the engine has been 
granted an exemption pursuant to the applicable provision of Sec. 34.7 
(a)(1), (a)(4), (b), (c), or (d), and an indication of the type of 
exemption and the reason for the grant must be noted in the permanent 
powerplant record that accompanies the engine from the time of 
manufacture of the engine.
    (iii) The designation NON-US indicates that the engine has been 
granted an exemption pursuant to Sec. 34.7(a)(1), and the notation 
``This aircraft may not be operated within the United States'', or an 
equivalent notation approved by the FAA, must be inserted in the 
aircraft logbook, or alternate equivalent document, at the time of 
installation of the engine.
    (iv) The designation EXEMPT NEW indicates that the engine has been 
granted an exemption pursuant to the applicable provision of Sec. 
34.7(h) of this chapter; the designation must be noted in the permanent 
powerplant record that accompanies the engine from the time of its 
manufacture.
    (v) The designation EXCEPTED SPARE indicates that the engine has 
been excepted pursuant to the applicable provision of Sec. 34.9(b) of 
this chapter; the designation must be noted in the permanent powerplant 
record that accompanies the engine from the time of its manufacture.
    (8) Any other information the FAA finds appropriate.
    (b) Except as provided in paragraph (d)(1) of this section, no 
person may remove, change, or place identification information required 
by paragraph (a) of this section, on any aircraft, aircraft engine, 
propeller, propeller blade, or propeller hub, without the approval of 
the FAA.
    (c) Except as provided in paragraph (d)(2) of this section, no 
person may remove or install any identification plate required by Sec. 
45.11, without the approval of the FAA.
    (d) Persons performing work under the provisions of Part 43 of this 
chapter may, in accordance with methods, techniques, and practices 
acceptable to the FAA--
    (1) Remove, change, or place the identification information required 
by paragraph (a) of this section on any aircraft, aircraft engine, 
propeller, propeller blade, or propeller hub; or
    (2) Remove an identification plate required by Sec. 45.11 when 
necessary during maintenance operations.
    (e) No person may install an identification plate removed in 
accordance with paragraph (d)(2) of this section on any aircraft, 
aircraft engine, propeller, propeller blade, or propeller hub other

[[Page 1034]]

than the one from which it was removed.

[Amdt. 45-3, 32 FR 188, Jan. 10, 1967, as amended by Amdt. 45-10, 44 FR 
45379, Aug. 2, 1979; Amdt. 45-12, 45 FR 60183, Sept. 11, 1980; Amdt. 45-
20, 55 FR 32861, Aug. 10, 1990; 55 FR 37287, Sept. 10, 1990; Amdt. 45-
26, 74 FR 53395, Oct. 16, 2009; Amdt. 45-28, 77 FR 76854, Dec. 31, 2012]



Sec. 45.15  Marking requirements for PMA articles, TSO articles, 
and Critical parts.

    (a) PMA articles. The manufacturer of a PMA article must permanently 
and legibly mark--
    (1) Each PMA article, with the PMA holder's name, trademark, symbol, 
or other FAA approved identification and part number; and
    (2) The letters ``FAA-PMA''.
    (b) TSO articles. The manufacturer of a TSO article must permanently 
and legibly mark --
    (1) Each TSO article with the TSO holder's name, trademark, symbol, 
or other FAA approved identification and part number; and
    (2) Each TSO article, unless otherwise specified in the applicable 
TSO, with the TSO number and letter of designation, all markings 
specifically required by the applicable TSO, and the serial number or 
the date of manufacture of the article or both.
    (c) Critical parts. Each person who manufactures a part for which a 
replacement time, inspection interval, or related procedure is specified 
in the Airworthiness Limitations section of a manufacturer's maintenance 
manual or Instructions for Continued Airworthiness must permanently and 
legibly mark that part with a serial number (or equivalent) unique to 
that part in addition to the other applicable requirements of this 
section.
    (d) If the FAA finds a part or article is too small or otherwise 
impractical to mark with any of the information required by this part, 
the manufacturer must attach that information to the part or its 
container.

[Doc. No. FAA-2006-25877, 74 FR 53395, Oct. 16, 2009]



Sec. 45.16  Marking of life-limited parts.

    When requested by a person required to comply with Sec. 43.10 of 
this chapter, the holder of a type certificate or design approval for a 
life-limited part must provide marking instructions, or must state that 
the part cannot be practicably marked without compromising its 
integrity.

[Doc. No. FAA-200-8017, 67 FR 2110, Jan. 15, 2002, as amended by Amdt. 
45-26, 74 FR 53395, Oct. 16, 2009]



              Subpart C_Nationality and Registration Marks



Sec. 45.21  General.

    (a) Except as provided in Sec. 45.22, no person may operate a U.S.-
registered aircraft unless that aircraft displays nationality and 
registration marks in accordance with the requirements of this section 
and Sec. Sec. 45.23 through 45.33.
    (b) Unless otherwise authorized by the FAA, no person may place on 
any aircraft a design, mark, or symbol that modifies or confuses the 
nationality and registration marks.
    (c) Aircraft nationality and registration marks must--
    (1) Except as provided in paragraph (d) of this section, be painted 
on the aircraft or affixed by any other means insuring a similar degree 
of permanence;
    (2) Have no ornamentation;
    (3) Contrast in color with the background; and
    (4) Be legible.
    (d) The aircraft nationality and registration marks may be affixed 
to an aircraft with readily removable material if--
    (1) It is intended for immediate delivery to a foreign purchaser;
    (2) It is bearing a temporary registration number; or
    (3) It is marked temporarily to meet the requirements of Sec. 
45.22(c)(1) or Sec. 45.29(h) of this part, or both.

[Doc. No. 8093, Amdt. 45-5, 33 FR 450, Jan. 12, 1968, as amended by 
Amdt. 45-17, 52 FR 34102, Sept. 9, 1987]



Sec. 45.22  Exhibition, antique, and other aircraft: Special rules.

    (a) When display of aircraft nationality and registration marks in 
accordance with Sec. Sec. 45.21 and 45.23 through 45.33 would be 
inconsistent with exhibition

[[Page 1035]]

of that aircraft, a U.S.-registered aircraft may be operated without 
displaying those marks anywhere on the aircraft if:
    (1) It is operated for the purpose of exhibition, including a motion 
picture or television production, or an airshow;
    (2) Except for practice and test fights necessary for exhibition 
purposes, it is operated only at the location of the exhibition, between 
the exhibition locations, and between those locations and the base of 
operations of the aircraft; and
    (3) For each flight in the United States:
    (i) It is operated with the prior approval of the Flight Standards 
District Office, in the case of a flight within the lateral boundaries 
of the surface areas of Class B, Class C, Class D, or Class E airspace 
designated for the takeoff airport, or within 4.4 nautical miles of that 
airport if it is within Class G airspace; or
    (ii) It is operated under a flight plan filed under either Sec. 
91.153 or Sec. 91.169 of this chapter describing the marks it displays, 
in the case of any other flight.
    (b) A small U.S.-registered aircraft built at least 30 years ago or 
a U.S.-registered aircraft for which an experimental certificate has 
been issued under Sec. 21.191(d) or 21.191(g) for operation as an 
exhibition aircraft or as an amateur-built aircraft and which has the 
same external configuration as an aircraft built at least 30 years ago 
may be operated without displaying marks in accordance with Sec. Sec. 
45.21 and 45.23 through 45.33 if:
    (1) It displays in accordance with Sec. 45.21(c) marks at least 2 
inches high on each side of the fuselage or vertical tail surface 
consisting of the Roman capital letter ``N'' followed by:
    (i) The U.S. registration number of the aircraft; or
    (ii) The symbol appropriate to the airworthiness certificate of the 
aircraft (``C'', standard; ``R'', restricted; ``L'', limited; or ``X'', 
experimental) followed by the U.S. registration number of the aircraft; 
and
    (2) It displays no other mark that begins with the letter ``N'' 
anywhere on the aircraft, unless it is the same mark that is displayed 
under paragraph (b)(1) of this section.
    (c) No person may operate an aircraft under paragraph (a) or (b) of 
this section--
    (1) In an ADIZ or DEWIZ described in Part 99 of this chapter unless 
it temporarily bears marks in accordance with Sec. Sec. 45.21 and 45.23 
through 45.33;
    (2) In a foreign country unless that country consents to that 
operation; or
    (3) In any operation conducted under Part 121, 133, 135, or 137 of 
this chapter.
    (d) If, due to the configuration of an aircraft, it is impossible 
for a person to mark it in accordance with Sec. Sec. 45.21 and 45.23 
through 45.33, he may apply to the FAA for a different marking 
procedure.

[Doc. No. 8093, Amdt. 45-5, 33 FR 450, Jan. 12, 1968, as amended by 
Amdt. 45-13, 46 FR 48603, Oct. 1, 1981; Amdt. 45-19, 54 FR 39291, Sept. 
25, 1989; Amdt. 45-18, 54 FR 34330, Aug. 18, 1989; Amdt. 45-21, 56 FR 
65653, Dec. 17, 1991; Amdt. 45-22, 66 FR 21066, Apr. 27, 2001]



Sec. 45.23  Display of marks; general.

    (a) Each operator of an aircraft must display on that aircraft marks 
consisting of the Roman capital letter ``N'' (denoting United States 
registration) followed by the registration number of the aircraft. Each 
suffix letter used in the marks displayed must also be a Roman capital 
letter.
    (b) When marks include only the Roman capital letter ``N'' and the 
registration number is displayed on limited, restricted or light-sport 
category aircraft or experimental or provisionally certificated 
aircraft, the operator must also display on that aircraft near each 
entrance to the cabin, cockpit, or pilot station, in letters not less 
than 2 inches nor more than 6 inches high, the words ``limited,'' 
``restricted,'' ``light-sport,'' ``experimental,'' or ``provisional,'' 
as applicable.

[Doc. No. 8093, Amdt. 45-5, 33 FR 450, Jan. 12, 1968, as amended by 
Amdt. 45-9, 42 FR 41102, Aug. 15, 1977; Amdt. 45-24, 69 FR 44863, July 
27, 2004]



Sec. 45.25  Location of marks on fixed-wing aircraft.

    (a) The operator of a fixed-wing aircraft must display the required 
marks on either the vertical tail surfaces or the sides of the fuselage, 
except as provided in Sec. 45.29(f).

[[Page 1036]]

    (b) The marks required by paragraph (a) of this section must be 
displayed as follows:
    (1) If displayed on the vertical tail surfaces, horizontally on both 
surfaces, horizontally on both surfaces of a single vertical tail or on 
the outer surfaces of a multivertical tail. However, on aircraft on 
which marks at least 3 inches high may be displayed in accordance with 
Sec. 45.29(b)(1), the marks may be displayed vertically on the vertical 
tail surfaces.
    (2) If displayed on the fuselage surfaces, horizontally on both 
sides of the fuselage between the trailing edge of the wing and the 
leading edge of the horizontal stabilizer. However, if engine pods or 
other appurtenances are located in this area and are an integral part of 
the fuselage side surfaces, the operator may place the marks on those 
pods or appurtenances.

[Amdt. 45-9, 42 FR 41102, Aug. 15, 1977]



Sec. 45.27  Location of marks; nonfixed-wing aircraft.

    (a) Rotorcraft. Each operator of a rotorcraft must display on that 
rotorcraft horizontally on both surfaces of the cabin, fuselage, boom, 
or tail the marks required by Sec. 45.23.
    (b) Airships. Each operator of an airship must display on that 
airship the marks required by Sec. 45.23, horizontally on--
    (1) The upper surface of the right horizontal stabilizer and on the 
under surface of the left horizontal stabilizer with the top of the 
marks toward the leading edge of each stabilizer; and
    (2) Each side of the bottom half of the vertical stabilizer.
    (c) Spherical balloons. Each operator of a spherical balloon must 
display the marks required by Sec. 45.23 in two places diametrically 
opposite and near the maximum horizontal circumference of that balloon.
    (d) Nonspherical balloons. Each operator of a nonspherical balloon 
must display the marks required by Sec. 45.23 on each side of the 
balloon near its maximum cross section and immediately above either the 
rigging band or the points of attachment of the basket or cabin 
suspension cables.
    (e) Powered parachutes and weight-shift-control aircraft. Each 
operator of a powered parachute or a weight-shift-control aircraft must 
display the marks required by Sec. Sec. 45.23 and 45.29(b)(2) of this 
part. The marks must be displayed in two diametrically opposite 
positions on the fuselage, a structural member, or a component of the 
aircraft and must be visible from the side of the aircraft.

[Doc. No. 2047, 29 FR 3223, Mar. 11, 1964, as amended by Amdt. 45-15, 48 
FR 11392, Mar. 17, 1983; Amdt. 45-24, 69 FR 44863, July 27, 2004; Amdt. 
45-25, 72 FR 52469, Sept. 14, 2007]



Sec. 45.29  Size of marks.

    (a) Except as provided in paragraph (f) of this section, each 
operator of an aircraft must display marks on the aircraft meeting the 
size requirements of this section.
    (b) Height. Except as provided in paragraph (h) of this part, the 
nationality and registration marks must be of equal height and on--
    (1) Fixed-wing aircraft, must be at least 12 inches high, except 
that:
    (i) An aircraft displaying marks at least 2 inches high before 
November 1, 1981 and an aircraft manufactured after November 2, 1981, 
but before January 1, 1983, may display those marks until the aircraft 
is repainted or the marks are repainted, restored, or changed;
    (ii) Marks at least 3 inches high may be displayed on a glider;
    (iii) Marks at least 3 inches high may be displayed on an aircraft 
for which the FAA has issued an experimental certificate under Sec. 
21.191 (d), Sec. 21.191 (g), or Sec. 21.191 (i) of this chapter to 
operate as an exhibition aircraft, an amateur-built aircraft, or a 
light-sport aircraft when the maximum cruising speed of the aircraft 
does not exceed 180 knots CAS; and
    (iv) Marks may be displayed on an exhibition, antique, or other 
aircraft in accordance with Sec. 45.22.
    (2) Airships, spherical balloons, nonspherical balloons, powered 
parachutes, and weight-shift-control aircraft must be at least 3 inches 
high; and
    (3) Rotorcraft, must be at least 12 inches high, except that 
rotorcraft displaying before April 18, 1983, marks required by Sec. 
45.29(b)(3) in effect on April 17, 1983, and rotorcraft manufactured on 
or after April 18, 1983, but before December 31, 1983, may display those

[[Page 1037]]

marks until the aircraft is repainted or the marks are repainted, 
restored, or changed.
    (c) Width. Characters must be two-thirds as wide as they are high, 
except the number ``1'', which must be one-sixth as wide as it is high, 
and the letters ``M'' and ``W'' which may be as wide as they are high.
    (d) Thickness. Characters must be formed by solid lines one-sixth as 
thick as the character is high.
    (e) Spacing. The space between each character may not be less than 
one-fourth of the character width.
    (f) If either one of the surfaces authorized for displaying required 
marks under Sec. 45.25 is large enough for display of marks meeting the 
size requirements of this section and the other is not, full size marks 
must be placed on the larger surface. If neither surface is large enough 
for full-size marks, marks as large as practicable must be displayed on 
the larger of the two surfaces. If no surface authorized to be marked by 
Sec. 45.27 is large enough for full-size marks, marks as large as 
practicable must be placed on the largest of the authorized surfaces. 
However, powered parachutes and weight-shift-control aircraft must 
display marks at least 3 inches high.
    (g) Uniformity. The marks required by this part for fixed-wing 
aircraft must have the same height, width, thickness, and spacing on 
both sides of the aircraft.
    (h) After March 7, 1988, each operator of an aircraft penetrating an 
ADIZ or DEWIZ must display on that aircraft temporary or permanent 
nationality and registration marks at least 12 inches high.

[Doc. No. 2047, 29 FR 3223, Mar. 11, 1964, as amended by Amdt. 45-2, 31 
FR 9863, July 21, 1966; Amdt. 45-9, 42 FR 41102, Aug. 15, 1977; Amdt. 
45-13, 46 FR 48604, Oct. 1, 1981; Amdt. 45-15, 48 FR 11392, Mar. 17, 
1983; Amdt. 45-17, 52 FR 34102, Sept. 9, 1987; 52 FR 36566, Sept. 30, 
1987; Amdt. 45-24, 69 FR 44863, July 27, 2004; Amdt.45-25, 72 FR 52469, 
Sept. 14, 2007]



Sec. 45.31  Marking of export aircraft.

    A person who manufactures an aircraft in the United States for 
delivery outside thereof may display on that aircraft any marks required 
by the State of registry of the aircraft. However, no person may operate 
an aircraft so marked within the United States, except for test and 
demonstration flights for a limited period of time, or while in 
necessary transit to the purchaser.



Sec. 45.33  Sale of aircraft; removal of marks.

    When an aircraft that is registered in the United States is sold, 
the holder of the Certificate of Aircraft Registration must remove, 
before its delivery to the purchaser, all United States marks from the 
aircraft, unless the purchaser is--
    (a) A citizen of the United States;
    (b) An individual citizen of a foreign country who is lawfully 
admitted for permanent residence in the United States; or
    (c) When the aircraft is to be based and primarily used in the 
United States, a corporation (other than a corporation which is a 
citizen of the United States) lawfully organized and doing business 
under the laws of the United States or any State thereof.

[Amdt. 45-11, 44 FR 61938, Oct. 29, 1979]



PART 47_AIRCRAFT REGISTRATION--Table of Contents



                            Subpart A_General

Sec.
47.1 Applicability.
47.2 Definitions.
47.3 Registration required.
47.5 Applicants.
47.7 United States citizens and resident aliens.
47.8 Voting trusts.
47.9 Corporations not U.S. citizens.
47.11 Evidence of ownership.
47.13 Signatures and instruments made by representatives.
47.15 Registration number.
47.16 Temporary registration numbers.
47.17 Fees.
47.19 Registry.

             Subpart B_Certificates of Aircraft Registration

47.31 Application.
47.33 Aircraft not previously registered anywhere.
47.35 Aircraft last previously registered in the United States.
47.37 Aircraft last previously registered in a foreign country.
47.39 Effective date of registration.

[[Page 1038]]

47.40 Registration expiration and renewal.
47.41 Duration and return of Certificate.
47.43 Invalid registration.
47.45 Change of address.
47.47 Cancellation of Certificate for export purpose.
47.49 Replacement of Certificate.
47.51 [Reserved]

          Subpart C_Dealers' Aircraft Registration Certificate

47.61 Dealer's Aircraft Registration Certificates.
47.63 Application.
47.65 Eligibility.
47.67 Evidence of ownership.
47.69 Limitations.
47.71 Duration of Certificate; change of status.

    Authority: 4 U.S.T. 1830; Pub. L. 108-297, 118 Stat. 1095 (49 U.S.C. 
40101 note, 49 U.S.C. 44101 note); 49 U.S.C. 106(g), 40113-40114, 44101-
44108, 44110-44113, 44703-44704, 44713, 45302, 46104, 46301.

    Source: Docket No. 7190, 31 FR 4495, Mar. 17, 1966, unless otherwise 
noted.

    Editorial Note: Nomenclature changes to part 47 appear at 75 FR 
41979, July 20, 2010.



                            Subpart A_General



Sec. 47.1  Applicability.

    This part prescribes the requirements for registering aircraft under 
49 U.S.C. 44101-44104. Subpart B applies to each applicant for, and 
holder of, a Certificate of Aircraft Registration, AC Form8050-3. 
Subpart C applies to each applicant for, and holder of, a Dealer's 
Aircraft Registration Certificate, AC Form 8050-6.

[Amdt. 47-29, 75 FR 41979, July 20, 2010]



Sec. 47.2  Definitions.

    The following are definitions of terms used in this part:
    Registry means the FAA, Civil Aviation Registry, Aircraft 
Registration Branch.
    Resident alien means an individual citizen of a foreign country 
lawfully admitted for permanent residence in the United States as an 
immigrant in conformity with the regulations of the Immigration and 
Naturalization Service of the Department of Justice (8 CFR Chapter 1).
    U.S. citizen means one of the following:
    (1) An individual who is a citizen of the United States or one of 
its possessions.
    (2) A partnership each of whose partners is an individual who is a 
citizen of the United States.
    (3) A corporation or association organized under the laws of the 
United States or a State, the District of Columbia, or a territory or 
possession of the United States, of which the president and at least 
two-thirds of the board of directors and other managing officers are 
citizens of the United States, which is under the actual control of 
citizens of the United States, and in which at least 75 percent of the 
voting interest is owned or controlled by persons that are citizens of 
the United States.

[Amdt. 47-20, 44 FR 61939, Oct. 29, 1979; Amdt. 47-27, 70 FR 244, Jan. 
3, 2005; Amdt. 47-29, 75 FR 41979, July 20, 2010]



Sec. 47.3  Registration required.

    (a) An aircraft may be registered under 49 U.S.C. 44103 only when 
the aircraft is not registered under the laws of a foreign country and 
is--
    (1) Owned by a citizen of the United States;
    (2) Owned by an individual citizen of a foreign country lawfully 
admitted for permanent residence in the United States;
    (3) Owned by a corporation not a citizen of the United States when 
the corporation is organized and doing business under the laws of the 
United States or a State within the United States, and the aircraft is 
based and primarily used in the United States; or
    (4) An aircraft of--
    (i) The United States Government; or
    (ii) A State, the District of Columbia, a territory or possession of 
the United States, or a political subdivision of a State, territory, or 
possession.
    (b) No person may operate an aircraft that is eligible for 
registration under 49 U.S.C. 44101-44104, unless the aircraft--
    (1) Has been registered by its owner;
    (2) Is carrying aboard the temporary authorization required by Sec. 
47.31(c); or
    (3) Is an aircraft of the Armed Forces.

[[Page 1039]]

    (c) Governmental units are those named in paragraph (a) of this 
section and Puerto Rico.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966, as amended by Amdt. 47-20, 44 
FR 61939, Oct. 29, 1979; Amdt. 47-27, 70 FR 244, Jan. 3, 2005; Amdt. 47-
29, 75 FR 41979, July 20, 2010]



Sec. 47.5  Applicants.

    (a) A person who wishes to register an aircraft in the United States 
must submit an Aircraft Registration Application, AC Form 8050-1 under 
this part.
    (b) An aircraft may be registered only by and in the legal name of 
its owner.
    (c) 49 U.S.C. 44103(c), provides that registration is not evidence 
of ownership of aircraft in any proceeding in which ownership by a 
particular person is in issue. The FAA does not issue any certificate of 
ownership or endorse any information with respect to ownership on a 
Certificate of Aircraft Registration, AC Form 8050-3. The FAA issues a 
Certificate of Aircraft Registration, AC Form 8050-3 to the person who 
appears to be the owner on the basis of the evidence of ownership 
submitted pursuant to Sec. 47.11 with the Aircraft Registration 
Application, or recorded at the Registry.
    (d) In this part, ``owner'' includes a buyer in possession, a 
bailee, or a lessee of an aircraft under a contract of conditional sale, 
and the assignee of that person.

[Amdt. 47-20, 44 FR 61939, Oct. 29, 1979, as amended by Amdt. 47-27, 70 
FR 244, Jan. 3, 2005; Amdt. 47-29, 75 FR 41979, July 20, 2010]



Sec. 47.7  United States citizens and resident aliens.

    (a) U.S. citizens. An applicant for aircraft registration under this 
part who is a U.S. citizen must certify to this in the Aircraft 
Registration Application, AC Form 8050-1.
    (b) Resident aliens. An applicant for aircraft registration under 49 
U.S.C. 44102 who is a resident alien must furnish a representation of 
permanent residence and the applicant's alien registration number issued 
by the Immigration and Naturalization Service.
    (c) Trustees. An applicant for aircraft registration under 49 U.S.C. 
44102 that holds legal title to an aircraft in trust must comply with 
the following requirements:
    (1) Each trustee must be either a U.S. citizen or a resident alien.
    (2) The applicant must submit with the Aircraft Registration 
Application--
    (i) A copy of each document legally affecting a relationship under 
the trust;
    (ii) If each beneficiary under the trust, including each person 
whose security interest in the aircraft is incorporated in the trust, is 
either a U.S. citizen or a resident alien, an affidavit by the applicant 
to that effect; and
    (iii) If any beneficiary under the trust, including any person whose 
security interest in the aircraft is incorporated in the trust, is not a 
U.S. citizen or resident alien, an affidavit from each trustee stating 
that the trustee is not aware of any reason, situation, or relationship 
(involving beneficiaries or other persons who are not U.S. citizens or 
resident aliens) as a result of which those persons together would have 
more than 25 percent of the aggregate power to influence or limit the 
exercise of the trustee's authority.
    (3) If persons who are neither U.S. citizens nor resident aliens 
have the power to direct or remove a trustee, either directly or 
indirectly through the control of another person, the trust instrument 
must provide that those persons together may not have more than 25 
percent of the aggregate power to direct or remove a trustee. Nothing in 
this paragraph prevents those persons from having more than 25 percent 
of the beneficial interest in the trust.
    (d) Partnerships. A partnership may apply for a Certificate of 
Aircraft Registration, AC Form 8050-3, under 49 U.S.C. 44102 only if 
each partner, whether a general or limited partner, is an individual who 
is a citizen of the United States. Nothing in this section makes 
ineligible for registration an aircraft which is not owned as a 
partnership asset but is co-owned by--
    (1) Resident aliens; or
    (2) One or more resident aliens and one or more U.S. citizens.

[Amdt. 47-20, 44 FR 61939, Oct. 29, 1979, as amended by Amdt. 47-27, 70 
FR 244, Jan. 3, 2005; Amdt. 47-29, 75 FR 41980, July 20, 2010]

[[Page 1040]]



Sec. 47.8  Voting trusts.

    (a) If a voting trust is used to qualify a domestic corporation as a 
U.S. citizen, the corporate applicant must submit to the Registry--
    (1) A true copy of the fully executed voting trust agreement, which 
must identify each voting interest of the applicant, and which must be 
binding upon each voting trustee, the applicant corporation, all foreign 
stockholders, and each other party to the transaction; and
    (2) An affidavit executed by each person designated as voting 
trustee in the voting trust agreement, in which each affiant 
represents--
    (i) That each voting trustee is a citizen of the United States 
within the meaning of 49 U.S.C. 40102(a)(15).
    (ii) That each voting trustee is not a past, present, or prospective 
director, officer, employee, attorney, or agent of any other party to 
the trust agreement;
    (iii) That each voting trustee is not a present or prospective 
beneficiary, creditor, debtor, supplier or contractor of any other party 
to the trust agreement;
    (iv) That each voting trustee is not aware of any reason, situation, 
or relationship under which any other party to the agreement might 
influence the exercise of the voting trustee's totally independent 
judgment under the voting trust agreement.
    (b) Each voting trust agreement submitted under paragraph (a)(1) of 
this section must provide for the succession of a voting trustee in the 
event of death, disability, resignation, termination of citizenship, or 
any other event leading to the replacement of any voting trustee. Upon 
succession, the replacement voting trustee shall immediately submit to 
the Registry the affidavit required by paragraph (a)(2) of this section.
    (c) If the voting trust terminates or is modified, and the result is 
less than 75 percent control of the voting interest in the corporation 
by citizens of the United States, a loss of citizenship of the holder of 
the Certificate of Aircraft Registration, AC Form 8050-3 occurs, and 
Sec. 47.41(a)(3) of this part applies.
    (d) A voting trust agreement may not empower a trustee to act 
through a proxy.

[Amdt. 47-20, 44 FR 61939, Oct. 29, 1979, as amended by Amdt. 47-27, 70 
FR 245, Jan. 3, 2005; Amdt. 47-29, 75 FR 41980, July 20, 2010]



Sec. 47.9  Corporations not U.S. citizens.

    (a) Each corporation applying for registration of an aircraft under 
49 U.S.C. 44102 must submit to the Registry with the Aircraft 
Registration Application, AC Form 8050-1--
    (1) A certified copy of its certificate of incorporation;
    (2) A certification that it is lawfully qualified to do business in 
one or more States;
    (3) A certification that the aircraft will be based and primarily 
used in the United States; and
    (4) The location where the records required by paragraph (e) of this 
section will be maintained.
    (b) For the purposes of registration, an aircraft is based and 
primarily used in the United States if the flight hours accumulated 
within the United States amount to at least 60 percent of the total 
flight hours of the aircraft during--
    (1) For aircraft registered on or before January 1, 1980, the 6-
calendar month period beginning on January 1, 1980, and each 6-calendar 
month period thereafter; and
    (2) For aircraft registered after January 1, 1980, the period 
consisting in the remainder of the registration month and the succeeding 
6 calendar months and each 6-calendar month period thereafter.
    (c) For the purpose of this section, only those flight hours 
accumulated during non-stop (except for stops in emergencies or for 
purposes of refueling) flight between two points in the United States, 
even if the aircraft is outside of the United States during part of the 
flight, are considered flight hours accumulated within the United 
States.
    (d) In determining compliance with this section, any periods during 
which the aircraft is not validly registered in the United States are 
disregarded.
    (e) The corporation that registers an aircraft pursuant to 49 U.S.C. 
44102 shall maintain, and make available for

[[Page 1041]]

inspection by the FAA upon request, records containing the total flight 
hours in the United States of the aircraft for three calendar years 
after the year in which the flight hours were accumulated.
    (f) The corporation that registers an aircraft pursuant to 49 U.S.C. 
44102 shall send to the Registry, at the end of each period of time 
described in paragraphs (b)(1) and (2) of this section, either--
    (1) A signed report containing--
    (i) The total time in service of the airframe as provided in Sec. 
91.417(a)(2)(i), accumulated during that period; and
    (ii) The total flight hours in the United States of the aircraft 
accumulated during that period; or
    (2) A signed statement that the total flight hours of the aircraft, 
while registered in the United States during that period, have been 
exclusively within the United States.

[Amdt. 47-20, 44 FR 61940, Oct. 29, 1979, as amended by Amdt. 47-24, 54 
FR 34330, Aug. 18, 1989; Amdt. 47-27, 70 FR 245, Jan. 3, 2005; Amdt. 47-
29, 75 FR 41979, July 20, 2010]



Sec. 47.11  Evidence of ownership.

    Except as provided in Sec. Sec. 47.33 and 47.35, each person that 
submits an Aircraft Registration Application, AC Form 8050-1 under this 
part must also submit the required evidence of ownership, recordable 
under Sec. Sec. 49.13 and 49.17 of this chapter, as follows:
    (a) The buyer in possession, the bailee, or the lessee of an 
aircraft under a contract of conditional sale must submit the contract. 
The assignee under a contract of conditional sale must submit both the 
contract (unless it is already recorded at the Registry), and his 
assignment from the original buyer, bailee, lessee, or prior assignee.
    (b) The repossessor of an aircraft must submit--
    (1) A Certificate of Repossession of Encumbered Aircraft, FAA Form 
8050-4, or its equivalent, signed by the applicant and stating that the 
aircraft was repossessed or otherwise seized under the security 
agreement involved and applicable local law;
    (2) The security agreement (unless it is already recorded at the 
Registry), or a copy thereof certified as true under Sec. 49.21 of this 
chapter; and
    (3) When repossession was through foreclosure proceedings resulting 
in sale, a bill of sale signed by the sheriff, auctioneer, or other 
authorized person who conducted the sale, and stating that the sale was 
made under applicable local law.
    (c) The buyer of an aircraft at a judicial sale, or at a sale to 
satisfy a lien or charge, must submit a bill of sale signed by the 
sheriff, auctioneer, or other authorized person who conducted the sale, 
and stating that the sale was made under applicable local law.
    (d) The owner of an aircraft, the title to which has been in 
controversy and has been determined by a court, must submit a certified 
copy of the decision of the court.
    (e) The executor or administrator of the estate of the deceased 
former owner of an aircraft must submit a certified copy of the letters 
testimentary or letters of administration appointing him executor or 
administrator. The Certificate of Aircraft Registration, AC Form 8050-3 
is issued to the applicant as executor or administrator.
    (f) The buyer of an aircraft from the estate of a deceased former 
owner must submit both a bill of sale, signed for the estate by the 
executor or administrator, and a certified copy of the letters 
testimentary or letters of administration. When no executor or 
administrator has been or is to be appointed, the applicant must submit 
both a bill of sale, signed by the heir-at-law of the deceased former 
owner, and an affidavit of the heir-at-law stating that no application 
for appointment of an executor or administrator has been made, that so 
far as he can determine none will be made, and that he is the person 
entitled to, or having the right to dispose of, the aircraft under 
applicable local law.
    (g) The guardian of another person's property that includes an 
aircraft must submit a certified copy of the order of the court 
appointing him guardian. The Certificate of Aircraft Registration is 
issued to the applicant as guardian.
    (h) The trustee of property that includes an aircraft, as described 
in Sec. 47.7(c), must submit either a certified copy of the order of 
the court appointing the trustee, or a complete and true

[[Page 1042]]

copy of the instrument creating the trust. If there is more than one 
trustee, each trustee must sign the Aircraft Registration Application. 
The Certificate of Aircraft Registration is issued to a single applicant 
as trustee, or to several trustees jointly as co-trustees.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966, as amended by Amdt. 47-20, 44 
FR 61940, Oct. 29, 1979; Amdt. 47-23, 53 FR 1915, Jan. 25, 1988; Amdt. 
47-29, 75 FR 41980, July 20, 2010]



Sec. 47.13  Signatures and instruments made by representatives.

    (a) Each person signing an Aircraft Registration Application, AC 
Form 8050-1, or a document submitted as supporting evidence under this 
part, must sign in ink or by other means acceptable to the FAA. If 
signed in ink, the Aircraft Registration Application must also have the 
typed or legibly printed name of each signer in the signature block.
    (b) When one or more persons doing business under a trade name 
submits an Aircraft Registration Application, a document submitted as 
supporting evidence under this part, or a request for cancellation of a 
Certificate of Aircraft Registration, AC Form 8050-3, the application, 
document, or request must be signed by, or on behalf of, each person who 
shares title to the aircraft.
    (c) When an agent submits an Aircraft Registration Application, a 
document submitted as supporting evidence under this part, or a request 
for cancellation of a Certificate of Aircraft Registration, on behalf of 
the owner, that agent must--
    (1) State the name of the owner on the application, document, or 
request;
    (2) Sign as agent or attorney-in-fact on the application, document, 
or request; and
    (3) Submit a signed power of attorney, or a true copy thereof 
certified under Sec. 49.21 of this chapter, with the application, 
document, or request.
    (d) When a corporation submits an Aircraft Registration Application, 
a document submitted as supporting evidence under this part, or a 
request for cancellation of a Certificate of Aircraft Registration, it 
must--
    (1) Have an authorized person sign, by means acceptable to the FAA, 
the application, document, or request;
    (2) Show the title of the signer's office on the application, 
document, or request; and
    (3) Submit a copy of the authorization from the board of directors 
to sign for the corporation, certified as true under Sec. 49.21 of this 
chapter by a corporate officer or other person in a managerial position 
therein, with the application, document, or request, unless--
    (i) The signer of the application, document, or request is a 
corporate officer or other person in a managerial position in the 
corporation and the title of his office is stated in connection with his 
signature; or
    (ii) A valid authorization to sign is on file at the Registry.
    (4) The provisions of paragraph (d)(3) of this section do not apply 
to an irrevocable deregistration and export request authorization when 
an irrevocable deregistration and export request authorization under the 
Cape Town Treaty is signed by a corporate officer and is filed with the 
Registry.
    (e) When a partnership submits an Aircraft Registration Application, 
a document submitted as supporting evidence under this part, or a 
request for cancellation of a Certificate of Aircraft Registration, it 
must--
    (1) State the full name of the partnership on the application, 
document, or request;
    (2) State the name of each general partner on the application, 
document, or request; and
    (3) Have a general partner sign the application, document, or 
request.
    (f) When co-owners, who are not engaged in business as partners, 
submit an Aircraft Registration Application, a document submitted as 
supporting evidence under this part, or a request for cancellation of a 
Certificate of Aircraft Registration, each person who shares title to 
the aircraft under the arrangement must sign the application, document, 
or request.
    (g) A power of attorney or other evidence of a person's authority to 
sign for another, submitted under this part, is valid for the purposes 
of this section, unless sooner revoked, until--
    (1) Its expiration date stated therein; or

[[Page 1043]]

    (2) If an expiration date is not stated therein, for not more than 3 
years after the date--
    (i) It is signed; or
    (ii) The grantor (a corporate officer or other person in a 
managerial position therein, where the grantor is a corporation) 
certifies in writing that the authority to sign shown by the power of 
attorney or other evidence is still in effect.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966, as amended by Amdt. 47-2, 31 
FR 15349, Dec. 8, 1966; Amdt. 47-3, 32 FR 6554, Apr. 28, 1967; Amdt. 47-
12, 36 FR 8661, May 11, 1971; Amdt. 47-27, 70 FR 245, Jan. 3, 2005; 
Amdt. 47-29, 75 FR 41980, July 20, 2010]



Sec. 47.15  Registration number.

    (a) Number required. An applicant for aircraft registration must 
place a U.S. registration number (registration mark) on the Aircraft 
Registration Application, AC Form 8050-1, and on any evidence submitted 
with the application. There is no charge for the assignment of numbers 
provided in this paragraph. This paragraph does not apply to an aircraft 
manufacturer who applies for a group of U.S. registration numbers under 
paragraph (c) of this section; a person who applies for a special 
registration number under paragraphs (d) through (f) of this section; or 
a holder of a Dealer's Aircraft Registration Certificate, AC Form 8050-
6, who applies for a temporary registration number under Sec. 47.16.
    (1) Aircraft not previously registered anywhere. The applicant must 
obtain the U.S. registration number from the Registry by request in 
writing describing the aircraft by make, type, model, and serial number 
(or, if it is amateur-built, as provided in Sec. 47.33(b)) and stating 
that the aircraft has not previously been registered anywhere. If the 
aircraft was brought into the United States from a foreign country, the 
applicant must submit evidence that the aircraft has never been 
registered in a foreign country.
    (2) Aircraft last previously registered in the United States. Unless 
the applicant applies for a different number under paragraphs (d) 
through (f) of this section, the applicant must place the U.S. 
registration number that is already assigned to the aircraft on the 
Aircraft Registration Application, and the supporting evidence. If there 
is no number assigned, the applicant must obtain a U.S. registration 
number from the Registry by making a written request that describes the 
aircraft by make, model, and serial number.
    (3) Aircraft last previously registered in a foreign country. 
Whether or not the foreign registration has ended, the applicant must 
obtain a U.S. registration number from the Registry for an aircraft last 
previously registered in a foreign country, by request in writing 
describing the aircraft by make, model, and serial number, accompanied 
by--
    (i) Evidence of termination of foreign registration in accordance 
with Sec. 47.37(b) or the applicant's affidavit showing that foreign 
registration has ended; or
    (ii) If foreign registration has not ended, the applicant's 
affidavit stating that the number will not be placed on the aircraft 
until foreign registration has ended.
    (4) Duration of a U.S. registration number assignment. Authority to 
use the registration number obtained under paragraph (a)(1), (2), or (3) 
of this section expires 90 days after the date it is issued unless the 
applicant submits an Aircraft Registration Application and complies with 
Sec. 47.33 or Sec. 47.37, as applicable, within that period of time. 
However, the applicant may obtain an extension of this 90-day period 
from the Registry if the applicant shows that the delay in complying 
with that section is due to circumstances beyond the applicant's 
control.
    (b) A U.S. registration number may not exceed five symbols in 
addition to the prefix letter ``N''. These symbols may be all numbers 
(N10000), one to four numbers and one suffix letter (N 1000A), or one to 
three numbers and two suffix letters (N 100AB). The letters ``I'' and 
``O'' may not be used. The first zero in a number must always be 
preceded by at least one of the numbers 1 through 9.
    (c) An aircraft manufacturer may apply to the Registry for enough 
U.S. registration numbers to supply estimated production for the next 18 
months. There is no charge for this allocation of numbers.

[[Page 1044]]

    (d) Any available, unassigned U.S. registration number may be 
assigned as a special registration number. An applicant who wants a 
special registration number or wants to change the registration number 
of his aircraft may apply for it to the Registry. The fee required by 
Sec. 47.17 must accompany the application.
    (e) [Reserved]
    (f) The Registry authorizes a special registration number change on 
the Assignment of Special Registration Numbers, AC Form 8050-64. The 
authorization expires one year from the date the Registry issues an 
Assignment of Special Registration Numbers unless the special 
registration number is permanently placed on the aircraft. Within five 
days after the special registration number is placed on the aircraft, 
the owner must complete and sign the Assignment of Special Registration 
Numbers, state the date the number was placed on the aircraft, and 
return the original form to the Registry. The duplicate of the 
Assignment of Special Registration Numbers and the present Certificate 
of Aircraft Registration, AC Form 8050-3, must be carried in the 
aircraft as temporary authority to operate it. This temporary authority 
is valid until the date the owner receives the revised Certificate of 
Aircraft Registration showing the new registration number, but in no 
case is it valid for more than 120 days from the date the number is 
placed on the aircraft.
    (g) [Reserved]
    (h) A special registration number may be reserved for no more than 1 
year. If a person wishes to renew his reservation from year to year, he 
must apply to the Registry for renewal and submit the fee required by 
Sec. 47.17 for a special registration number.
    (i) When aircraft registration has ended, as described in Sec. 
47.41(a), the assignment of a registration number to an aircraft is no 
longer authorized for use except as provided in Sec. 47.31(c) and will 
be cancelled:
    (1) Following the date established in Sec. 47.40(a)(1) for any 
aircraft that has not been re-registered under Sec. 47.40(a);
    (2) Following the expiration date shown on the Certificate of 
Aircraft Registration for any aircraft whose registration has not been 
renewed under Sec. 47.40(c);
    (3) Following the expiration date shown on the Dealer's Aircraft 
Registration Certificate, AC Form 8050-6, for any aircraft registered 
under Subpart C of this part, when the certificate has not been renewed, 
and the owner has not applied for registration in accordance with Sec. 
47.31; or
    (4) When ownership has transferred--
    (i) Six months after first receipt of notice of aircraft sale or 
evidence of ownership from the last registered owner or successive 
owners, and an Aircraft Registration Application has not been received.
    (ii) Six months after evidence of ownership authorized under Sec. 
47.67 has been submitted, and the applicant has not met the requirements 
of this part.
    (iii) Twelve months after a new owner has submitted evidence of 
ownership and an Aircraft Registration Application under Sec. 47.31, 
and the applicant or a successive applicant has not met the requirements 
of this part.
    (j) At the time an assignment of registration number is cancelled, 
the number may be reserved for one year in the name of the last owner of 
record if a request has been submitted with the fee required by Sec. 
47.17. If the request for reservation and fee are not submitted prior to 
cancellation, the registration number is unavailable for assignment for 
a period of five years.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966, as amended by Amdt. 47-1, 31 
FR 13314, Oct. 14, 1966; Amdt. 47-5, 32 FR 13505, Sept. 27, 1967; Amdt. 
47-7, 34 FR 2480, Feb. 21, 1969; Amdt. 47-13, 36 FR 16187, Aug. 20, 
1971; Amdt. 47-15, 37 FR 21528, Oct. 12, 1972; Amdt. 47-16, 37 FR 25487, 
Dec. 1, 1972; Amdt. 47-17, 39 FR 1353, Jan. 8, 1974; Amdt. 47-22, 47 FR 
12153, Mar. 22, 1982; Amdt. 47-29, 75 FR 41980, July 20, 2010]



Sec. 47.16  Temporary registration numbers.

    (a) Temporary registration numbers are issued by the FAA to 
manufacturers, distributors, and dealers who are holders of Dealer's 
Aircraft Registration Certificates, AC Form 8050-6, for temporary 
display on aircraft during flight allowed under Subpart C of this part.
    (b) The holder of a Dealer's Aircraft Registration Certificate may 
apply to the Registry for as many temporary

[[Page 1045]]

registration numbers as are necessary for his business. The application 
must be in writing and include--
    (1) Sufficient information to justify the need for the temporary 
registration numbers requested; and
    (2) The number of each Dealer's Aircraft Registration Certificate 
held by the applicant.

There is no charge for these numbers.
    (c) The use of temporary registration numbers is subject to the 
following conditions:
    (1) The numbers may be used and reused--
    (i) Only in connection with the holder's Dealer's Aircraft 
Registration Certificate;
    (ii) Within the limitations of Sec. 47.69 where applicable, 
including the requirements of Sec. 47.67; and
    (iii) On aircraft not registered under Subpart B of this part or in 
a foreign country, and not displaying any other identification markings.
    (2) A temporary registration number may not be used on more than one 
aircraft in flight at the same time.
    (3) Temporary registration numbers may not be used to fly aircraft 
into the United States for the purpose of importation.
    (d) The assignment of any temporary registration number to any 
person lapses upon the expiration of all of his Dealer's Aircraft 
Registration Certificates. When a temporary registration number is used 
on a flight outside the United States for delivery purposes, the holder 
shall record the assignment of that number to the aircraft and shall 
keep that record for at least 1 year after the removal of the number 
from that aircraft. Whenever the owner of an aircraft bearing a 
temporary registration number applies for an airworthiness certificate 
under Part 21 of this chapter he shall furnish that number in the 
application. The temporary registration number must be removed from the 
aircraft not later than the date on which either title or possession 
passes to another person.

[Amdt. 47-4, 32 FR 12556, Aug. 30, 1967, as amended by Amdt. 47-29, 75 
FR 41981, July 20, 2010]



Sec. 47.17  Fees.

    (a) The fees for applications under this part are as follows:

(1) Certificate of Aircraft Registration (each aircraft)....       $5.00
(2) Dealer's Aircraft Registration Certificate..............       10.00
(3) Additional Dealer's Aircraft Registration Certificate           2.00
 (issued to same dealer)....................................
(4) Special registration number (each number)...............       10.00
(5) To change, reassign, or reserve a registration number...       10.00
(6) Replacement Certificate of Aircraft Registration........        2.00
(7) Re-registration or Renewal Certificate of Aircraft              5.00
 Registration...............................................
 

    (b) Each application must be accompanied by the proper fee, that may 
be paid by check or money order to the Federal Aviation Administration.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966; 31 FR 5483, Apr. 7, 1966, as 
amended by Doc. No. 8084, 32 FR 5769, Apr. 11, 1967; Amdt. 47-29, 75 FR 
41981, July 20, 2010]



Sec. 47.19  Registry.

    Each application, request, notification, or other communication sent 
to the FAA under this part must be mailed to the Registry, Department of 
Transportation, Post Office Box 25504, Oklahoma City, Oklahoma 73125-
0504, or delivered to the Registry at 6425 S. Denning Ave., Oklahoma 
City, Oklahoma 73169.

[Amdt. 47-27, 70 FR 245, Jan. 3, 2005]



             Subpart B_Certificates of Aircraft Registration



Sec. 47.31  Application.

    (a) Each applicant for a Certificate of Aircraft Registration, AC 
Form 8050-3 must submit the following to the Registry--
    (1) An Aircraft Registration Application, AC Form 8050-1, signed by 
the applicant in the manner prescribed by Sec. 47.13;
    (2) The original Aircraft Bill of Sale, AC Form 8050-2, or other 
evidence of ownership authorized by Sec. 47.33, Sec. 47.35, or Sec. 
47.37 (unless already recorded at the Registry); and
    (3) The fee required by Sec. 47.17.
    (b) The FAA rejects an application when--
    (1) Any form is not completed;
    (2) The name and signature of the applicant are not the same 
throughout; or

[[Page 1046]]

    (3) The applicant does not provide a legibly printed or typed name 
with the signature in the signature block.
    (c) After compliance with paragraph (a) of this section, the 
applicant for registration of an aircraft last previously registered in 
the United States must carry the second copy of the Aircraft 
Registration Application in the aircraft as temporary authority to 
operate without registration.
    (1) This temporary authority is valid for operation within the 
United States until the date the applicant receives the Certificate of 
Aircraft Registration or until the date the FAA denies the application, 
but in no case for more than 90 days after the date the applicant signs 
the application. If by 90 days after the date the applicant signs the 
Aircraft Registration Application, the FAA has neither issued the 
Certificate of Aircraft Registration nor denied the application, the 
Registry will issue a letter of extension that serves as authority to 
continue to operate the aircraft without registration while it is 
carried in the aircraft.
    (2) This temporary authority is not available in connection with any 
Aircraft Registration Application received when 12 months have passed 
since the receipt of the first application following transfer of 
ownership by the last registered owner.
    (3) If there is no registration number assigned at the time 
application for registration is made, the second copy of the Aircraft 
Registration Application may not be used as temporary authority to 
operate the aircraft.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966; 31 FR 5483, Apr. 7, 1966, as 
amended by Amdt. 47-6, 33 FR 11, Jan. 3, 1968; Amdt. 47-15, 37 FR 21528, 
Oct. 12, 1972; Amdt. 47-16, 37 FR 25487, Dec. 1, 1972; Amdt. 47-28, 73 
FR 10667, Feb. 28, 2008; Amdt. 47-29, 75 FR 41981, July 20, 2010]



Sec. 47.33  Aircraft not previously registered anywhere.

    (a) A person who is the owner of an aircraft that has not been 
registered under 49 U.S.C. 44101-44104, under other law of the United 
States, or under foreign law, may register it under this part if he--
    (1) Complies with Sec. Sec. 47.3, 47.7, 47.8, 47.9, 47.11, 47.13, 
47.15, and 47.17, as applicable; and
    (2) Submits with his Aircraft Registration Application, AC Form 
8050-1, an Aircraft Bill of Sale, AC Form 8050-2, signed by the seller, 
an equivalent bill of sale, or other evidence of ownership authorized by 
Sec. 47.11.
    (b) If, for good reason, the applicant cannot produce the evidence 
of ownership required by paragraph (a) of this section, he must submit 
other evidence that is satisfactory to the FAA. This other evidence may 
be an affidavit stating why he cannot produce the required evidence, 
accompanied by whatever further evidence is available to prove the 
transaction.
    (c) The owner of an amateur-built aircraft who applies for 
registration under paragraphs (a) and (b) of this section must describe 
the aircraft by class (airplane, rotorcraft, glider, or balloon), serial 
number, number of seats, type of engine installed, (reciprocating, 
turbopropeller, turbojet, or other), number of engines installed, and 
make, model, and serial number of each engine installed; and must state 
whether the aircraft is built for land or water operation. Also, he must 
submit as evidence of ownership an affidavit giving the U.S. 
registration number, and stating that the aircraft was built from parts 
and that he is the owner. If he built the aircraft from a kit, the 
applicant must also submit a bill of sale from the manufacturer of the 
kit.
    (d) The owner, other than the holder of the type certificate, of an 
aircraft that he assembles from parts to conform to the approved type 
design, must describe the aircraft and engine in the manner required by 
paragraph (c) of this section, and also submit evidence of ownership 
satisfactory to the FAA, such as bills of sale, for all major components 
of the aircraft.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966; 31 FR 5483, Apr. 7, 1966, as 
amended by Amdt. 47-16, 37 FR 25487, Dec. 1, 1972; Amdt. 47-20, 44 FR 
61940, Oct. 29, 1979; Amdt. 47-27, 70 FR 245, Jan. 3, 2005; Amdt. 47-29, 
75 FR 41979, July 20, 2010]



Sec. 47.35  Aircraft last previously registered in the United States.

    (a) A person who is the owner of an aircraft last previously 
registered under 49 U.S.C. Sections 44101-44104, or under other law of 
the United States,

[[Page 1047]]

may register it under this part if he complies with Sec. Sec. 47.3, 
47.7, 47.8, 47.9, 47.11, 47.13, 47.15, and 47.17, as applicable and 
submits with his Aircraft Registration Application, AC Form 8050-1 an 
Aircraft Bill of Sale, AC Form 8050-2, signed by the seller or an 
equivalent conveyance, or other evidence of ownership authorized by 
Sec. 47.11.
    (1) If the applicant bought the aircraft from the last registered 
owner, the conveyance must be from that owner to the applicant.
    (2) If the applicant did not buy the aircraft from the last 
registered owner, he must submit conveyances or other instruments 
showing consecutive transactions from the last registered owner through 
each intervening owner to the applicant.
    (b) If, for good reason, the applicant cannot produce the evidence 
of ownership required by paragraph (a) of this section, he must submit 
other evidence that is satisfactory to the FAA. This other evidence may 
be an affidavit stating why he cannot produce the required evidence, 
accompanied by whatever further evidence is available to prove the 
transaction.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966, as amended by Amdt. 47-16, 37 
FR 25487, Dec. 1, 1972; Amdt. 47-20, 44 FR 61940, Oct. 29, 1979; Amdt. 
47-27, 70 FR 245, Jan. 3, 2005; 73 FR 55722, Sept. 26, 2008; Amdt. 47-
29, 75 FR 41979, July 20, 2010]



Sec. 47.37  Aircraft last previously registered in a foreign country.

    (a) A person who is the owner of an aircraft last previously 
registered under the law of a foreign country may register it under this 
part if the owner--
    (1) Complies with Sec. Sec. 47.3, 47.7, 47.8, 47.9, 47.11, 47.13, 
47.15, and 47.17, as applicable;
    (2) Submits with his Aircraft Registration Application, AC Form 
8050-1 a bill of sale from the foreign seller or other evidence 
satisfactory to the FAA that he owns the aircraft; and
    (3) Submits evidence satisfactory to the FAA that--
    (i) If the country in which the aircraft was registered has not 
ratified the Convention on the International Recognition of Rights in 
Aircraft (4 U.S.T. 1830), (the Geneva Convention), or the Convention on 
International Interests in Mobile Equipment, as modified by the Protocol 
to the Convention on International Interests in Mobile Equipment on 
Matters Specific to Aircraft Equipment (the Cape Town Treaty), the 
foreign registration has ended or is invalid; or
    (ii) If that country has ratified the Geneva Convention, but has not 
ratified the Cape Town Treaty, the foreign registration has ended or is 
invalid, and each holder of a recorded right against the aircraft has 
been satisfied or has consented to the transfer, or ownership in the 
country of export has been ended by a sale in execution under the terms 
of the Geneva Convention; or
    (iii) If that country has ratified the Cape Town Treaty and the 
aircraft is subject to the Treaty, that the foreign registration has 
ended or is invalid, and that all interests ranking in priority have 
been discharged or that the holders of such interests have consented to 
the deregistration and export of the aircraft.
    (iv) Nothing under (a)(3)(iii) affects rights established prior to 
the Treaty entering into force with respect to the country in which the 
aircraft was registered.
    (b) For the purposes of paragraph (a)(3) of this section, 
satisfactory evidence of termination of the foreign registration may 
be--
    (1) A statement, by the official having jurisdiction over the 
national aircraft registry of the foreign country, that the registration 
has ended or is invalid, and showing the official's name and title and 
describing the aircraft by make, model, and serial number; or
    (2) A final judgment or decree of a court of competent jurisdiction 
of the foreign country, determining that, under the laws of that 
country, the registration has become invalid.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966, as amended by Amdt. 47-20, 44 
FR 61940, Oct. 29, 1979; Amdt. 47-26, 68 FR 10317, Mar. 4, 2003; Amdt. 
47-27, 70 FR 245, Jan. 3, 2005]



Sec. 47.39  Effective date of registration.

    An aircraft is registered on the date the Registry determines that 
the submissions meet the requirements of this part. The effective date 
of registration

[[Page 1048]]

is shown by a date stamp on the Aircraft Registration Application, AC 
Form 8050-1, and as the date of issue on the Certificate of Aircraft 
Registration, AC Form 8050-3.

[Amdt. 47-29, 75 FR 41981, July 20, 2010]



Sec. 47.40  Registration expiration and renewal.

    (a) Re-registration. Each aircraft registered under this part before 
October 1, 2010, must be re-registered in accordance with this paragraph 
(a).
    (1) A Certificate of Aircraft Registration issued before October 1, 
2010, expires on the expiration date identified in the following 
schedule that corresponds with the month in which the certificate was 
issued.

------------------------------------------------------------------------
                                                        The owner must
                                                         apply for re-
                                                         registration
  If the certificate was issued     The certificate      between these
               in:                    expires on:       dates--to allow
                                                        delivery of the
                                                        new certificate
                                                       before expiration
------------------------------------------------------------------------
March of any year...............  March 31, 2011....  November 1, 2010
                                                       and January 31,
                                                       2011.
April of any year...............  June 30, 2011.....  February 1, 2011
                                                       and April 30,
                                                       2011.
May of any year.................  September 30, 2011  May 1, 2011 and
                                                       July 31, 2011.
June of any year................  December 31, 2011.  August 1, 2011 and
                                                       October 31, 2011.
July of any year................  March 31, 2012....  November 1, 2011
                                                       and January 31,
                                                       2012.
August of any year..............  June 30, 2012.....  February 1, 2012
                                                       and April 30,
                                                       2012.
September of any year...........  September 30, 2012  May 1, 2012 and
                                                       July 31, 2012.
October of any year.............  December 31, 2012.  August 1, 2012 and
                                                       October 31, 2012.
November of any year............  March 31 2013.....  November 1, 2012
                                                       and January 31,
                                                       2013.
December of any year............  June 30, 2013.....  February 1, 2013
                                                       and April 30,
                                                       2013.
January of any year.............  September 30, 2013  May 1, 2013 and
                                                       July 31, 2013.
February of any year............  December 31, 2013.  August 1, 2013 and
                                                       October 31, 2013.
------------------------------------------------------------------------

    (2) Each holder of a Certificate of Aircraft Registration, AC Form 
8050-3, issued before October 1, 2010, must submit an Application for 
Aircraft Re-registration, AC Form 8050-1A, and the fee required by Sec. 
47.17, between October 1, 2010, and December 31, 2013, according to the 
schedule in paragraph (a)(1) of this section.
    (3) A Certificate of Aircraft Registration issued under this 
paragraph expires three years after the last day of the month in which 
it is issued.
    (b) Initial Registration. A Certificate of Aircraft Registration 
issued in accordance with Sec. 47.31 expires three years after the last 
day of the month in which it is issued.
    (c) Renewal. Each holder of a Certificate of Aircraft Registration, 
AC Form 8050-3, containing an expiration date may apply for renewal by 
submitting an Application for Aircraft Registration Renewal, AC Form 
8050-1B, and the fee required by Sec. 47.17 during the six months 
preceding the expiration date. A certificate issued under this paragraph 
expires three years from the expiration date of the previous 
certificate.

[Amdt. 47-29, 75 FR 41981, July 20, 2010]



Sec. 47.41  Duration and return of Certificate.

    (a) Each Certificate of Aircraft Registration, AC Form 8050-3, 
issued by the FAA under this subpart is effective, unless registration 
has ended by reason of having been revoked, canceled, expired, or the 
ownership is transferred, until the date upon which one of the following 
events occurs:
    (1) Subject to the Convention on the International Recognition of 
Rights in Aircraft when applicable, the aircraft is registered under the 
laws of a foreign country.
    (2) The aircraft is totally destroyed or scrapped.
    (3) The holder of the certificate loses his U.S. citizenship.
    (4) 30 days have elapsed since the death of the holder of the 
certificate.
    (5) The owner, if an individual who is not a citizen of the United 
States, loses status as a resident alien, unless that person becomes a 
citizen of the United States at the same time.
    (6) If the owner is a corporation other than a corporation which is 
a citizen of the United States--
    (i) The corporation ceases to be lawfully organized and doing 
business under the laws of the United States or any State thereof; or

[[Page 1049]]

    (ii) A period described in Sec. 47.9(b) ends and the aircraft was 
not based and primarily used in the United States during that period.
    (7) If the trustee in whose name the aircraft is registered--
    (i) Loses U.S. citizenship;
    (ii) Loses status as a resident alien and does not become a citizen 
of the United States at the same time; or
    (iii) In any manner ceases to act as trustee and is not immediately 
replaced by another who meets the requirements of Sec. 47.7(c).
    (b) The Certificate of Aircraft Registration, with the reverse side 
completed, must be returned to the Registry--
    (1) Within 21 days in the case of registration under the laws of a 
foreign country, by the person who was the owner of the aircraft before 
foreign registration;
    (2) Within 60 days after the death of the holder of the certificate, 
by the administrator or executor of his estate, or by his heir-at-law if 
no administrator or executor has been or is to be appointed; or
    (3) Within 21 days of the termination of the registration, by the 
holder of the Certificate of Aircraft Registration in all other cases 
mentioned in paragraph (a) of this section, except in the case of 
expired certificates, the holder must destroy the expired certificate.
    (4) If the certificate is not available for return, as directed in 
paragraph (b) of this section, a statement describing the aircraft and 
stating the reason the certificate is not available must be submitted to 
the Registry within the time required by paragraph (b) of this section.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966; 31 FR 5483, Apr. 7, 1966, as 
amended by Amdt. 47-20, 44 FR 61940, Oct. 29, 1979; Amdt. 47-28, 73 FR 
10667, Feb. 28, 2008; Amdt. 47-29, 75 FR 41982, July 20, 2010]



Sec. 47.43  Invalid registration.

    (a) The registration of an aircraft is invalid if, at the time it is 
made--
    (1) The aircraft is registered in a foreign country;
    (2) The applicant is not the owner;
    (3) The applicant is not qualified to submit an application under 
this part; or
    (4) The interest of the applicant in the aircraft was created by a 
transaction that was not entered into in good faith, but rather was made 
to avoid (with or without the owner's knowledge) compliance with 49 
U.S.C. 44101-44104.
    (b) If the registration of an aircraft is invalid under paragraph 
(a) of this section, the holder of the invalid Certificate of Aircraft 
Registration, AC Form 8050-3, must return it as soon as possible to the 
Registry.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966; 31 FR 5483, Apr. 7, 1966, as 
amended by Amdt. 47-20, 44 FR 61940, Oct. 29, 1979; Amdt. 47-27, 70 FR 
245, Jan. 3, 2005; Amdt. 47-29, 75 FR 41982, July 20, 2010]



Sec. 47.45  Change of address.

    Within 30 days after any change in a registered owner's mailing 
address, the registered owner must notify the Registry in writing of the 
change of address. If a post office box or mailing drop is used for 
mailing purposes, the registered owner also must provide that owner's 
physical address or location. Upon acceptance, the Registry will issue, 
without charge, a revised Certificate of Aircraft Registration, AC Form 
8050-3, reflecting the new mailing address. When a post office box or 
mailing drop is used for mailing purposes, and the registered owner's 
physical address or location changes, the registered owner must notify 
the Registry in writing of the new address or location within 30 days.

[Amdt. 47-29, 75 FR 41982, July 20, 2010]



Sec. 47.47  Cancellation of Certificate for export purpose.

    (a) The holder of a Certificate of Aircraft Registration, AC Form 
8050-3, or the holder of an irrevocable deregistration and export 
request authorization recognized under the Cape Town Treaty and filed 
with the FAA, who wishes to cancel the Certificate of Aircraft 
Registration for the purpose of export must submit to the Registry--
    (1) A written request for cancellation of the Certificate of 
Aircraft Registration describing the aircraft by make, model, and serial 
number, and stating the U.S. registration number and the country to 
which the aircraft will be exported;

[[Page 1050]]

    (2)(i) For an aircraft not subject to the Cape Town Treaty, evidence 
satisfactory to the FAA that each holder of a recorded right has been 
satisfied or has consented to the transfer; or
    (ii) For an aircraft subject to the Cape Town Treaty, evidence 
satisfactory to the FAA that each holder of a recorded right established 
prior to the date the Treaty entered into force with respect to the 
United States has been satisfied or has consented to the transfer; and
    (3) A written certification that all registered interests ranking in 
priority to that of the requestor have been discharged or that the 
holders of such interests have consented to the cancellation for export 
purposes.
    (b) If the aircraft is subject to the Cape Town Treaty and an 
irrevocable deregistration and export request authorization has been 
filed with the Registry, the Registry will honor a request for 
cancellation only if an authorized party makes the request.
    (c) The Registry notifies the country to which the aircraft is to be 
exported of the cancellation.

[Amdt. 47-27, 70 FR 245, Jan. 3, 2005, as amended by Amdt. 47-29, 75 FR 
41982, July 20, 2010]



Sec. 47.49  Replacement of Certificate.

    (a) If the original Certificate of Aircraft Registration, AC Form 
8050-3, is lost, stolen, or mutilated, the registered owner may submit 
to the Registry a written request that states the reason a replacement 
certificate is needed and the fee required by Sec. 47.17. The Registry 
will send a replacement certificate to the registered owner's mailing 
address or to another mailing address if requested in writing by the 
registered owner.
    (b) The registered owner may request a temporary Certificate of 
Aircraft Registration pending receipt of a replacement certificate. The 
Registry issues a temporary Certificate of Aircraft Registration in the 
form of a fax that must be carried in the aircraft until receipt of the 
replacement certificate.

[Amdt. 47-29, 75 FR 41982, July 20, 2010]



Sec. 47.51  [Reserved]



          Subpart C_Dealers' Aircraft Registration Certificate



Sec. 47.61  Dealer's Aircraft Registration Certificates.

    (a) The FAA issues a Dealer's Aircraft Registration Certificate, AC 
Form 8050-6, to U.S. manufacturers and dealers to--
    (1) Allow manufacturers to make any required flight tests of 
aircraft.
    (2) Facilitate operating, demonstrating, and merchandising aircraft 
by the manufacturer or dealer without the burden of obtaining a 
Certificate of Aircraft Registration, AC Form 8050-3, for each aircraft 
with each transfer of ownership, under subpart B of this part.
    (b) A Dealer's Aircraft Registration Certificate is an alternative 
for the Certificate of Aircraft Registration issued under subpart B of 
this part. A dealer may, under this subpart, obtain one or more Dealer's 
Aircraft Registration Certificates in addition to his original 
certificate, and he may use a Dealer's Aircraft Registration Certificate 
for any aircraft he owns.
    (c) If the Dealer's Aircraft Registration Certificate expires under 
Sec. 47.71, and an aircraft is registered under this subpart, 
application for registration must be made under Sec. 47.31, or the 
assignment of registration number may be cancelled in accordance with 
Sec. 47.15(i)(3).

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966; as amended by Amdt. 47-9, 35 
FR 802, Jan. 21, 1970; Amdt. 47-16, 37 FR 25487, Dec. 1, 1972; Amdt. 47-
29, 75 FR 41982, July 20, 2010]



Sec. 47.63  Application.

    A manufacturer or dealer that wishes to obtain a Dealer's Aircraft 
Registration Certificate, AC Form 8050-6, must submit--
    (a) A Dealer's Aircraft Registration Certificate Application, AC 
Form 8050-5; and
    (b) The fee required by Sec. 47.17.

[Doc. No. 7190, 31 FR 4495, Mar. 17, 1966, as amended by Amdt. 47-16, 37 
FR 25487, Dec. 1, 1972; Amdt. 47-29, 75 FR 41982, July 20, 2010]

[[Page 1051]]



Sec. 47.65  Eligibility.

    To be eligible for a Dealer's Aircraft Registration Certificate, AC 
Form 8050-6, the applicant must have an established place of business in 
the United States, must be substantially engaged in manufacturing or 
selling aircraft, and must be a citizen of the United States, as defined 
by 49 U.S.C. 40102 (a)(15).

[Amdt. 47-29, 75 FR 41983, July 20, 2010]



Sec. 47.67  Evidence of ownership.

    Before using a Dealer's Aircraft Registration Certificate, AC Form 
8050-6, for operating the aircraft, the holder of the certificate (other 
than a manufacturer) must send to the Registry evidence of ownership 
under Sec. 47.11. An Aircraft Bill of Sale, AC Form 8050-2, or its 
equivalent, may be used as evidence of ownership. There is no recording 
fee.

[Amdt. 47-29, 75 FR 41983, July 20, 2010]



Sec. 47.69  Limitations.

    A Dealer's Aircraft Registration Certificate, AC Form 8050-6 is 
valid only in connection with use of aircraft--
    (a) By the owner of the aircraft to whom it was issued, his agent or 
employee, or a prospective buyer, and in the case of a dealer other than 
a manufacturer, only after he has complied with Sec. 47.67;
    (b) Within the United States, except when used to deliver to a 
foreign purchaser an aircraft displaying a temporary registration number 
and carrying an airworthiness certificate on which that number is 
written;
    (c) While a certificate is carried within the aircraft; and
    (d) On a flight that is--
    (1) For required flight testing of aircraft; or
    (2) Necessary for, or incident to, sale of the aircraft.

However, a prospective buyer may operate an aircraft for demonstration 
purposes only while he is under the direct supervision of the holder of 
the Dealer's Aircraft Registration Certificate or his agent.

[Doc. No. 7190 31 FR 4495, Mar. 17, 1966; 31 FR 5483, Apr. 7, 1966, as 
amended by Amdt. 47-4, 32 FR 12556, Aug. 30, 1967; Amdt. 47-29, 75 FR 
41983, July 20, 2010]



Sec. 47.71  Duration of Certificate; change of status.

    (a) A Dealer's Aircraft Registration Certificate, AC Form 8050-6, 
expires 1 year after the date it is issued. Each additional certificate 
expires on the date the original certificate expires.
    (b) The holder of a Dealer's Aircraft Registration Certificate must 
immediately notify the Registry of any of the following--
    (1) A change of name;
    (2) A change of address;
    (3) A change that affects status as a citizen of the United States; 
or
    (4) The discontinuance of business.

[31 FR 4495, Mar. 17, 1966, as amended by Amdt. 47-29, 75 FR 41983, July 
20, 2010]



PART 49_RECORDING OF AIRCRAFT TITLES AND SECURITY DOCUMENTS--
Table of Contents



                         Subpart A_Applicability

Sec.
49.1 Applicability.

                            Subpart B_General

49.11 FAA Aircraft Registry.
49.13 Signatures and acknowledgements.
49.15 Fees for recording.
49.17 Conveyances recorded.
49.19 Effective date of filing for recordation.
49.21 Return of original conveyance.

     Subpart C_Aircraft Ownership and Encumbrances Against Aircraft

49.31 Applicability.
49.33 Eligibility for recording: general requirements.
49.35 Eligibility for recording: ownership requirements.
49.37 Claims for salvage or extraordinary expenses.

Subpart D_Encumbrances Against Specifically Identified Aircraft Engines 
                             and Propellers

49.41 Applicability.
49.43 Eligibility for recording: general requirements.

[[Page 1052]]

49.45 Recording of releases, cancellations, discharges, and 
          satisfactions: special requirements.

Subpart E_Encumbrances Against Air Carrier Aircraft Engines, Propellers, 
                       Appliances, and Spare Parts

49.51 Applicability.
49.53 Eligibility for recording: general requirements.
49.55 Recording of releases, cancellations, discharges, and 
          satisfactions: special requirements.

   Subpart F_Transmission of Information to the International Registry

49.61 Applicability.
49.63 Eligibility for Authorization for Transmission to the 
          International Registry: General Requirements.

    Authority: 4 U.S.T. 1830; Pub. L. 108-297, 118 Stat. 1095 (49 U.S.C. 
40101 note, 49 U.S.C. 44101 note); 49 U.S.C. 106(g), 40113-40114, 44101-
44108, 44110-44113, 44704, 44713, 45302, 46104, 46301.

    Source: Docket No. 1996, 29 FR 6486, May 19, 1964, unless otherwise 
noted.



                         Subpart A_Applicability



Sec. 49.1  Applicability.

    (a) This part applies to the recording of certain conveyances 
affecting title to, or any interest in--
    (1) Any aircraft registered under 49 U.S.C. 44101-44104;
    (2) Any specifically identified aircraft engine of 750 or more rated 
takeoff horsepower, or the equivalent of that horsepower;
    (3) Any specifically identified aircraft propeller able to absorb 
750 or more rated takeoff shaft horsepower; and
    (4) Any aircraft engine, propeller, or appliance maintained by or 
for an air carrier certificated under 49 U.S.C. 44705, for installation 
or use in an aircraft, aircraft engine, or propeller, or any spare part, 
maintained at a designated location or locations by or for such an air 
carrier.
    (b) Subpart B of this part governs, where applicable by its terms, 
conveyances subject to this part.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-10, 70 
FR 246, Jan. 3, 2005]



                            Subpart B_General



Sec. 49.11  FAA Aircraft Registry.

    To be eligible for recording, a conveyance must be mailed to the FAA 
Aircraft Registry, Department of Transportation, Post Office Box 25504, 
Oklahoma City, Oklahoma 73125-0504, or delivered to the Registry at 6425 
S. Denning Ave., Oklahoma City, Oklahoma 73169.

[Amdt. 49-10, 70 FR 246, Jan. 3, 2005]



Sec. 49.13  Signatures and acknowledgements.

    (a) Each signature on a conveyance must be in ink.
    (b) Paragraphs (b) through (f) of Sec. 47.13 of this chapter apply 
to a conveyance made by, or on behalf of, one or more persons doing 
business under a trade name, or by an agent, corporation, partnership, 
coowner, or unincorporated association.
    (c) No conveyance or other instrument need be acknowledged, as 
provided in 49 U.S.C. 44107(c), in order to be recorded under this part. 
The law of the place of delivery of the conveyance determines when a 
conveyance or other instrument must be acknowledged in order to be valid 
for the purposes of that place.
    (d) A power of attorney or other evidence of a person's authority to 
sign for another, submitted under this part, is valid for the purposes 
of this section, unless sooner revoked, until--
    (1) Its expiration date stated therein; or
    (2) If an expiration date is not stated thereon, for not more than 3 
years after the date--
    (i) It is signed; or
    (ii) The grantor (a corporate officer or other person in a 
managerial position therein, where the grantor is a corporation) 
certifies in writing that the authority to sign shown by the power of 
attorney or other evidence is still in effect.

[Doc. No. 7190, 31 FR 4499, Mar. 17, 1966, as amended by Amdt. 49-2, 31 
FR 15349, Dec. 8, 1966; Amdt. 49-6, 36 FR 8661, May 11, 1971; Amdt. 49-
10, 70 FR 246, Jan. 3, 2005]

[[Page 1053]]



Sec. 49.15  Fees for recording.

    (a) The fees charged for recording conveyances under this part are 
as follows:

(1) Conveyance of aircraft--
  For each aircraft listed therein..........................       $5.00
(2) Conveyance, made for security purposes, of a
 specifically identified aircraft engine or propeller, or
 any assignment or amendment thereof, or supplement thereto,
 recorded under Subpart D--
  For each engine or propeller..............................        5.00
(3) Conveyance, made for security purposes, of aircraft
 engines, propellers, appliances, or spare parts, maintained
 at a designated location or locations, or any assignment or
 amendment thereof, or supplement thereto, recorded under
 Subpart E--
  For the group of items at each location...................        5.00
 

    (b) There is no fee for recording a bill of sale that accompanies an 
application for aircraft registration and the proper fee under Part 47 
of this chapter.
    (c) Each conveyance must be accompanied by the proper fee, that may 
be paid by check or money order to the Federal Aviation Administration.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-1, 31 
FR 4499, Mar. 17, 1966; Doc. No. 8084, 32 FR 5769, Apr. 11, 1967]



Sec. 49.17  Conveyances recorded.

    (a)(1) Each instrument recorded under this part is a ``conveyance'' 
within the following definition in 49 U.S.C. 40102(a)(19):

    ``Conveyance'' means an instrument, including a conditional sales 
contract, affecting title to, or an interest in, property.

    (2) A notice of Federal tax lien is not recordable under this part, 
since it is required to be filed elsewhere by the Internal Revenue Code 
(26 U.S.C. 6321, 6323; 26 CFR 301.6321-1, 301.6323-1).
    (b) The kinds of conveyance recordable under this part include those 
used as evidence of ownership under Sec. 47.11 of this chapter.
    (c) The validity of any instrument, eligible for recording under 
this part, is governed by the laws of the State, possession, Puerto 
Rico, or District of Columbia, as the case may be, in which the 
instrument was delivered, regardless of the location or place of 
delivery of the property affected by the instrument. If the place where 
an instrument is intended to be delivered is stated in the instrument, 
it is presumed that the instrument was delivered at that place. The 
recording of a conveyance is not a decision of the FAA that the 
instrument does, in fact, affect title to, or an interest in, the 
aircraft or other property it covers.
    (d) The following rules apply to conveyances executed for security 
purposes and assignments thereof:
    (1) A security agreement must be signed by the debtor. If the debtor 
is not the registered owner of the aircraft, the security agreement must 
be accompanied by the debtor's Application for Aircraft Registration and 
evidence of ownership, as prescribed in Part 47 of this chapter, unless 
the debtor--
    (i) Holds a Dealer's Aircraft Registration Certificate and submits 
evidence of ownership as provided in Sec. 47.67 of this chapter (if 
applicable);
    (ii) Was the owner of the aircraft on the date the security 
agreement was signed, as shown by documents recorded at the FAA Aircraft 
Registry; or
    (iii) Is the vendor, bailor, or lessor under a contract of 
conditional sale.
    (2) The name of a cosigner may not appear in the security agreement 
as a debtor or owner. If a person other than the registered owner signs 
the security agreement, that person must show the capacity in which that 
person signs, such as ``cosigner'' or ``guarantor''.
    (3) An assignment of an interest in a security agreement must be 
signed by the assignor and, unless it is attached to and is a part of 
the original agreement, must describe the agreement in sufficient detail 
to identify it, including its date, the names of the parties, the date 
of FAA recording, and the recorded conveyance number.
    (4) An amendment of, or a supplement to, a conveyance executed for 
security purposes that has been recorded by the FAA must meet the 
requirements for recording the original conveyance and must describe the 
original conveyance in sufficient detail to identify it, including its 
date, the names of the parties, the date of FAA recording, and the 
recorded conveyance number.
    (5) Immediately after a debt secured by a conveyance given for 
security purposes has been satisfied, or any of the encumbered aircraft 
have been released

[[Page 1054]]

from the conveyance, the holder shall execute a release on AC Form 8050-
41, Part II--Release, provided to him by the FAA when the conveyance was 
recorded by the FAA, or its equivalent, and shall send it to the FAA 
Aircraft Registry for recording. If the debt is secured by more than one 
aircraft and all of the collateral is released, the collateral need not 
be described in detail in the release. However, the original conveyance 
must be clearly described in enough detail to identify it, including its 
date, the names of the parties, the date of FAA recording, and the 
recorded conveyance number.
    (6) A contract of conditional sale, as defined in 49 U.S.C. 
40102(a)(18), must be signed by all parties to the contract.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-1, 31 
FR 4499, Mar. 17, 1966; Amdt. 49-9, 53 FR 1915, Jan. 25, 1988; Amdt. 49-
10, 70 FR 246, Jan. 3, 2005]



Sec. 49.19  Effective date of filing for recordation.

    A conveyance is filed for recordation upon the date and at the time 
it is received by the FAA Aircraft Registry.



Sec. 49.21  Return of original conveyance.

    If a person submitting a conveyance for recording wants the original 
returned to him, he must submit a true copy with the original. After 
recording, the copy is kept by the FAA and the original is returned to 
the applicant stamped with the date and time of recording. The copy must 
be imprinted on paper permanent in nature, including dates, and 
signatures, to which is attached a certificate of the person submitting 
the conveyance stating that the copy has been compared with the original 
and that it is a true copy.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-1, 31 
FR 4499, Mar. 17, 1966]



     Subpart C_Aircraft Ownership and Encumbrances Against Aircraft



Sec. 49.31  Applicability.

    This subpart applies to the recording of the following kinds of 
conveyances:
    (a) A bill of sale, contract of conditional sale, assignment of an 
interest under a contract of conditional sale, mortgage, assignment of 
mortgage, lease, equipment trust, notice of tax lien or of other lien, 
or other instrument affecting title to, or any interest in, aircraft.
    (b) A release, cancellation, discharge, or satisfaction of a 
conveyance named in paragraph (a) of this section.



Sec. 49.33  Eligibility for recording: general requirements.

    A conveyance is eligible for recording under this subpart only if, 
in addition to the requirements of Sec. Sec. 49.11, 49.13, and 49.17, 
the following requirements are met:
    (a) It is in a form prescribed by, or acceptable to, the 
Administrator for that kind of conveyance;
    (b) It describes the aircraft by make and model, manufacturer's 
serial number, and United States registration number, or other detail 
that makes identification possible;
    (c) It is an original document, or a duplicate original document, or 
if neither the original nor a duplicate original of a document is 
available, a true copy of an original document, certified under Sec. 
49.21;
    (d) It affects aircraft registered under 49 U.S.C. 44101-44104; and
    (e) It is accompanied by the recording fee required by Sec. 49.15, 
but there is no fee for recording a conveyance named in Sec. 49.31(b).

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-1, 31 
FR 4499, Mar. 17, 1966; Amdt. 49-10, 70 FR 246, Jan. 3, 2005]



Sec. 49.35  Eligibility for recording: ownership requirements.

    If the seller of an aircraft is not shown on the records of the FAA 
as the owner of the aircraft, a conveyance, including a contract of 
conditional sale, submitted for recording under this subpart must be 
accompanied by bills of sale or similar documents showing consecutive 
transfers from the last registered owner, through each intervening 
owner, to the seller.



Sec. 49.37  Claims for salvage or extraordinary expenses.

    The right to a charge arising out of a claim for compensation for 
salvage of

[[Page 1055]]

an aircraft or for extraordinary expenses indispensable for preserving 
the aircraft in operations terminated in a foreign country that is a 
party to the Convention on the International Recognition of Rights in 
Aircraft (4 U.S.T. 1830) may be noted on the FAA record by filing notice 
thereof with the FAA Aircraft Registry within three months after the 
date of termination of the salvage or preservation operations.



Subpart D_Encumbrances Against Specifically Identified Aircraft Engines 
                             and Propellers



Sec. 49.41  Applicability.

    This subpart applies to the recording of the following kinds of 
conveyances:
    (a) Any lease, a notice of tax lien or other lien (except a notice 
of Federal tax lien referred to in Sec. 49.17(a)), and any mortgage, 
equipment trust, contract of conditional sale, or other instrument 
executed for security purposes, which affects title to, or any interest 
in, any specifically identified aircraft engine of 550 or more rated 
takeoff horsepower, or the equivalent of that horsepower, or a 
specifically identified aircraft propeller capable of absorbing 750 or 
more rated takeoff shaft horsepower.
    (b) An assignment or amendment of, or supplement to, an instrument 
named in paragraph (a) of this section.
    (c) A release, cancellation, discharge, or satisfaction of a 
conveyance named in paragraph (a) or (b) of this section.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-5, 35 
FR 802, Jan. 21, 1970; Amdt. 49-10, 70 FR 246, Jan. 3, 2005]



Sec. 49.43  Eligibility for recording: general requirements.

    A conveyance is eligible for recording under this subpart only if, 
in addition to the requirements of Sec. Sec. 49.11, 49.13, and 49.17, 
the following requirements are met:
    (a) It affects and describes an aircraft engine or propeller to 
which this subpart applies, specifically identified by make, model, 
horsepower, and manufacturer's serial number; and
    (b) It is accompanied by the recording fee required by Sec. 49.15, 
but there is no fee for recording a conveyance named in Sec. 49.41(c).



Sec. 49.45  Recording of releases, cancellations, discharges,
and satisfactions: special requirements.

    (a) A release, cancellation, discharge, or satisfaction of an 
encumbrance created by an instrument recorded under this subpart must be 
in a form equivalent to AC Form 8050-41 and contain a description of the 
encumbrance, the recording information furnished to the holder at the 
time of recording, and the collateral released.
    (b) If more than one engine or propeller, or both, are listed in an 
instrument, recorded under this subpart, that created an encumbrance 
thereon and all of them are released, they need not be listed by serial 
number, but the release, cancellation, discharge, or satisfaction must 
state that all of the encumbered engines or propellers are released. The 
original recorded document must be clearly identified by the names of 
the parties, the date of FAA recording, and the document date.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-7, 37 
FR 25487, Dec. 1, 1972]



Subpart E_Encumbrances Against Air Carrier Aircraft Engines, Propellers, 
                       Appliances, and Spare Parts



Sec. 49.51  Applicability.

    This subpart applies to the recording of the following kinds of 
conveyances:
    (a) Any lease, a notice of tax lien or other lien (except a notice 
of Federal tax lien referred to in Sec. 49.17 (a), and any mortgage, 
equipment trust, contract of conditional sale, or other instrument 
executed for security purposes, which affects title to, or any interest 
in, any aircraft engine, propeller, or appliance maintained by or on 
behalf of an air carrier certificated under 49 U.S.C. 44705 for 
installation or use in aircraft, aircraft engines, or propellers, or any 
spare parts, maintained at a designated location or locations by or on 
behalf of such an air carrier.
    (b) An assignment or amendment of, or supplement to, an instrument 
named in paragraph (a) of this section.

[[Page 1056]]

    (c) A release, cancellation, discharge, or satisfaction of a 
conveyance named in paragraph (a) or (b) of this section.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-5, 35 
FR 802, Jan. 21, 1970; Amdt. 49-10, 70 FR 246, Jan. 3, 2005]



Sec. 49.53  Eligibility for recording: general requirements.

    (a) A conveyance is eligible for recording under this subpart only 
if, in addition to the requirements of Sec. Sec. 49.11, 49.13, and 
49.17, the following requirements are met:
    (1) It affects any aircraft engine, propeller, appliance, or spare 
part, maintained by or on behalf of an air carrier certificated under 49 
U.S.C. 44705;
    (2) It contains or is accompanied by a statement by the air carrier 
certificated under that section;
    (3) It specifically describes the location or locations of each 
aircraft engine, propeller, appliance, or spare part covered by it; and
    (4) It is accompanied by the recording fee required by Sec. 49.15, 
but there is no fee for recording a conveyance named in Sec. 49.51(c).
    (b) The conveyance need only describe generally, by type, the 
engines, propellers, appliances, or spare parts covered by it.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-5, 35 
FR 802, Jan. 21, 1970; Amdt. 49-10, 70 FR 246, Jan. 3, 2005]



Sec. 49.55  Recording of releases, cancellations, discharges,
and satisfactions: special requirements.

    (a) A release, cancellation, discharge, or satisfaction of an 
encumbrance on all of the collateral listed in an instrument recorded 
under this subpart, or on all of the collateral at a particular 
location, must be in a form equivalent to AC Form 8050-41, signed by the 
holder of all of the collateral at the particular location, and contain 
a description of the encumbrance, the recording information furnished to 
the holder at the time of recording, and the location of the released 
collateral.
    (b) If the encumbrance on collateral at all of the locations listed 
in an instrument recorded under this subpart is released, canceled, 
discharged, or satisfied, the locations need not be listed. However, the 
document must state that all of the collateral at all of the locations 
listed in the encumbrance has been so released, canceled, discharged, or 
satisfied. The original recorded document must be clearly identified by 
the names of the parties, the date of recording by the FAA, and the 
document number.

[Doc. No. 1996, 29 FR 6486, May 19, 1964, as amended by Amdt. 49-1, 31 
FR 4499, Mar. 17, 1966; Amdt. 49-7, 37 FR 25487, Dec. 1, 1972]



   Subpart F_Transmission of Information to the International Registry

    Source: Docket No. FAA-2004-19944, 70 FR 246, Jan. 3, 2005, unless 
otherwise noted.

    Effective Date Note: At 70 FR 247, Jan. 3, 2005, subpart F was 
added. This subpart contains information collection and recordkeeping 
requirements and will not become effective until approval has been given 
by the Office of Management and Budget.



Sec. 49.61  Applicability.

    The FAA Civil Aviation Registry is designated under Section 3 of the 
Cape Town Treaty Implementation Act of 2004, as the entry point for 
authorizing the transmission of information to the International 
Registry affecting United States civil aircraft, aircraft assigned a 
U.S. registration number and engines with a rated shaft horsepower of 
550 or the equivalent thereof. This subpart applies to the transmission 
of information to the International Registry; the filing of the Entry 
Point filing form, AC Form 8050-135; and the filing of documents 
eligible for recording under subparts C and D of part 49.



Sec. 49.63  Eligibility for Authorization for Transmission to the
International Registry: general requirements.

    (a) To send information to the International Registry with respect 
to a civil aircraft of the United States, an aircraft for which a U.S. 
identification number has been assigned, or an aircraft engine, a person 
requesting a unique authorization code from the FAA Aircraft Registry 
must comply with the following:
    (1) File a completed AC Form 8050-135 with the FAA Aircraft 
Registry; and

[[Page 1057]]

    (2) For civil aircraft of the United States, file with the FAA 
Aircraft Registry any documents representing the transaction that meet 
the requirements of subpart C of this part; or
    (3) For aircraft engines, file with the FAA Aircraft Registry any 
documents representing the transaction that meet the requirements of 
subpart D of this part.
    (b) Nothing in this section requires transmittal of information 
relating to aircraft engines to the International Registry through the 
FAA Aircraft Registry.

                         PARTS 50	59 [RESERVED]

[[Page 1059]]



                              FINDING AIDS




  --------------------------------------------------------------------

  A list of CFR titles, subtitles, chapters, subchapters and parts and 
an alphabetical list of agencies publishing in the CFR are included in 
the CFR Index and Finding Aids volume to the Code of Federal Regulations 
which is published separately and revised annually.

  Table of CFR Titles and Chapters
  Alphabetical List of Agencies Appearing in the CFR
  List of CFR Sections Affected

[[Page 1061]]



                    Table of CFR Titles and Chapters




                     (Revised as of January 1, 2015)

                      Title 1--General Provisions

         I  Administrative Committee of the Federal Register 
                (Parts 1--49)
        II  Office of the Federal Register (Parts 50--299)
       III  Administrative Conference of the United States (Parts 
                300--399)
        IV  Miscellaneous Agencies (Parts 400--500)

                    Title 2--Grants and Agreements

            Subtitle A--Office of Management and Budget Guidance 
                for Grants and Agreements
         I  Office of Management and Budget Governmentwide 
                Guidance for Grants and Agreements (Parts 2--199)
        II  Office of Management and Budget Guidance (Parts 200--
                299)
            Subtitle B--Federal Agency Regulations for Grants and 
                Agreements
       III  Department of Health and Human Services (Parts 300--
                399)
        IV  Department of Agriculture (Parts 400--499)
        VI  Department of State (Parts 600--699)
       VII  Agency for International Development (Parts 700--799)
      VIII  Department of Veterans Affairs (Parts 800--899)
        IX  Department of Energy (Parts 900--999)
        XI  Department of Defense (Parts 1100--1199)
       XII  Department of Transportation (Parts 1200--1299)
      XIII  Department of Commerce (Parts 1300--1399)
       XIV  Department of the Interior (Parts 1400--1499)
        XV  Environmental Protection Agency (Parts 1500--1599)
     XVIII  National Aeronautics and Space Administration (Parts 
                1800--1899)
        XX  United States Nuclear Regulatory Commission (Parts 
                2000--2099)
      XXII  Corporation for National and Community Service (Parts 
                2200--2299)
     XXIII  Social Security Administration (Parts 2300--2399)
      XXIV  Housing and Urban Development (Parts 2400--2499)
       XXV  National Science Foundation (Parts 2500--2599)
      XXVI  National Archives and Records Administration (Parts 
                2600--2699)
     XXVII  Small Business Administration (Parts 2700--2799)
    XXVIII  Department of Justice (Parts 2800--2899)

[[Page 1062]]

      XXIX  Department of Labor (Parts 2900--2999)
       XXX  Department of Homeland Security (Parts 3000--3099)
      XXXI  Institute of Museum and Library Services (Parts 3100--
                3199)
     XXXII  National Endowment for the Arts (Parts 3200--3299)
    XXXIII  National Endowment for the Humanities (Parts 3300--
                3399)
     XXXIV  Department of Education (Parts 3400--3499)
      XXXV  Export-Import Bank of the United States (Parts 3500--
                3599)
     XXXVI  Office of National Drug Control Policy, Executive 
                Office of the President (Parts 3600--3699)
    XXXVII  Peace Corps (Parts 3700--3799)
     LVIII  Election Assistance Commission (Parts 5800--5899)
       LIX  Gulf Coast Ecosystem Restoration Council (Parts 5900--
                5999)

                        Title 3--The President

         I  Executive Office of the President (Parts 100--199)

                           Title 4--Accounts

         I  Government Accountability Office (Parts 1--199)
        II  Recovery Accountability and Transparency Board (Parts 
                200--299)

                   Title 5--Administrative Personnel

         I  Office of Personnel Management (Parts 1--1199)
        II  Merit Systems Protection Board (Parts 1200--1299)
       III  Office of Management and Budget (Parts 1300--1399)
         V  The International Organizations Employees Loyalty 
                Board (Parts 1500--1599)
        VI  Federal Retirement Thrift Investment Board (Parts 
                1600--1699)
      VIII  Office of Special Counsel (Parts 1800--1899)
        IX  Appalachian Regional Commission (Parts 1900--1999)
        XI  Armed Forces Retirement Home (Parts 2100--2199)
       XIV  Federal Labor Relations Authority, General Counsel of 
                the Federal Labor Relations Authority and Federal 
                Service Impasses Panel (Parts 2400--2499)
        XV  Office of Administration, Executive Office of the 
                President (Parts 2500--2599)
       XVI  Office of Government Ethics (Parts 2600--2699)
       XXI  Department of the Treasury (Parts 3100--3199)
      XXII  Federal Deposit Insurance Corporation (Parts 3200--
                3299)
     XXIII  Department of Energy (Parts 3300--3399)
      XXIV  Federal Energy Regulatory Commission (Parts 3400--
                3499)
       XXV  Department of the Interior (Parts 3500--3599)
      XXVI  Department of Defense (Parts 3600--3699)
    XXVIII  Department of Justice (Parts 3800--3899)

[[Page 1063]]

      XXIX  Federal Communications Commission (Parts 3900--3999)
       XXX  Farm Credit System Insurance Corporation (Parts 4000--
                4099)
      XXXI  Farm Credit Administration (Parts 4100--4199)
    XXXIII  Overseas Private Investment Corporation (Parts 4300--
                4399)
     XXXIV  Securities and Exchange Commission (Parts 4400--4499)
      XXXV  Office of Personnel Management (Parts 4500--4599)
    XXXVII  Federal Election Commission (Parts 4700--4799)
        XL  Interstate Commerce Commission (Parts 5000--5099)
       XLI  Commodity Futures Trading Commission (Parts 5100--
                5199)
      XLII  Department of Labor (Parts 5200--5299)
     XLIII  National Science Foundation (Parts 5300--5399)
       XLV  Department of Health and Human Services (Parts 5500--
                5599)
      XLVI  Postal Rate Commission (Parts 5600--5699)
     XLVII  Federal Trade Commission (Parts 5700--5799)
    XLVIII  Nuclear Regulatory Commission (Parts 5800--5899)
      XLIX  Federal Labor Relations Authority (Parts 5900--5999)
         L  Department of Transportation (Parts 6000--6099)
       LII  Export-Import Bank of the United States (Parts 6200--
                6299)
      LIII  Department of Education (Parts 6300--6399)
       LIV  Environmental Protection Agency (Parts 6400--6499)
        LV  National Endowment for the Arts (Parts 6500--6599)
       LVI  National Endowment for the Humanities (Parts 6600--
                6699)
      LVII  General Services Administration (Parts 6700--6799)
     LVIII  Board of Governors of the Federal Reserve System 
                (Parts 6800--6899)
       LIX  National Aeronautics and Space Administration (Parts 
                6900--6999)
        LX  United States Postal Service (Parts 7000--7099)
       LXI  National Labor Relations Board (Parts 7100--7199)
      LXII  Equal Employment Opportunity Commission (Parts 7200--
                7299)
     LXIII  Inter-American Foundation (Parts 7300--7399)
      LXIV  Merit Systems Protection Board (Parts 7400--7499)
       LXV  Department of Housing and Urban Development (Parts 
                7500--7599)
      LXVI  National Archives and Records Administration (Parts 
                7600--7699)
     LXVII  Institute of Museum and Library Services (Parts 7700--
                7799)
    LXVIII  Commission on Civil Rights (Parts 7800--7899)
      LXIX  Tennessee Valley Authority (Parts 7900--7999)
       LXX  Court Services and Offender Supervision Agency for the 
                District of Columbia (Parts 8000--8099)
      LXXI  Consumer Product Safety Commission (Parts 8100--8199)
    LXXIII  Department of Agriculture (Parts 8300--8399)
     LXXIV  Federal Mine Safety and Health Review Commission 
                (Parts 8400--8499)
     LXXVI  Federal Retirement Thrift Investment Board (Parts 
                8600--8699)

[[Page 1064]]

    LXXVII  Office of Management and Budget (Parts 8700--8799)
      LXXX  Federal Housing Finance Agency (Parts 9000--9099)
   LXXXIII  Special Inspector General for Afghanistan 
                Reconstruction (Parts 9300--9399)
    LXXXIV  Bureau of Consumer Financial Protection (Parts 9400--
                9499)
    LXXXVI  National Credit Union Administration (Parts 9600--
                9699)
     XCVII  Department of Homeland Security Human Resources 
                Management System (Department of Homeland 
                Security--Office of Personnel Management) (Parts 
                9700--9799)
     XCVII  Council of the Inspectors General on Integrity and 
                Efficiency (Parts 9800--9899)
      XCIV  Military Compensation and Retirement Modernization 
                Commission (Parts 9900--9999)

                      Title 6--Domestic Security

         I  Department of Homeland Security, Office of the 
                Secretary (Parts 1--199)
         X  Privacy and Civil Liberties Oversight Board (Parts 
                1000--1099)

                         Title 7--Agriculture

            Subtitle A--Office of the Secretary of Agriculture 
                (Parts 0--26)
            Subtitle B--Regulations of the Department of 
                Agriculture
         I  Agricultural Marketing Service (Standards, 
                Inspections, Marketing Practices), Department of 
                Agriculture (Parts 27--209)
        II  Food and Nutrition Service, Department of Agriculture 
                (Parts 210--299)
       III  Animal and Plant Health Inspection Service, Department 
                of Agriculture (Parts 300--399)
        IV  Federal Crop Insurance Corporation, Department of 
                Agriculture (Parts 400--499)
         V  Agricultural Research Service, Department of 
                Agriculture (Parts 500--599)
        VI  Natural Resources Conservation Service, Department of 
                Agriculture (Parts 600--699)
       VII  Farm Service Agency, Department of Agriculture (Parts 
                700--799)
      VIII  Grain Inspection, Packers and Stockyards 
                Administration (Federal Grain Inspection Service), 
                Department of Agriculture (Parts 800--899)
        IX  Agricultural Marketing Service (Marketing Agreements 
                and Orders; Fruits, Vegetables, Nuts), Department 
                of Agriculture (Parts 900--999)
         X  Agricultural Marketing Service (Marketing Agreements 
                and Orders; Milk), Department of Agriculture 
                (Parts 1000--1199)
        XI  Agricultural Marketing Service (Marketing Agreements 
                and Orders; Miscellaneous Commodities), Department 
                of Agriculture (Parts 1200--1299)

[[Page 1065]]

       XIV  Commodity Credit Corporation, Department of 
                Agriculture (Parts 1400--1499)
        XV  Foreign Agricultural Service, Department of 
                Agriculture (Parts 1500--1599)
       XVI  Rural Telephone Bank, Department of Agriculture (Parts 
                1600--1699)
      XVII  Rural Utilities Service, Department of Agriculture 
                (Parts 1700--1799)
     XVIII  Rural Housing Service, Rural Business-Cooperative 
                Service, Rural Utilities Service, and Farm Service 
                Agency, Department of Agriculture (Parts 1800--
                2099)
        XX  Local Television Loan Guarantee Board (Parts 2200--
                2299)
       XXV  Office of Advocacy and Outreach, Department of 
                Agriculture (Parts 2500--2599)
      XXVI  Office of Inspector General, Department of Agriculture 
                (Parts 2600--2699)
     XXVII  Office of Information Resources Management, Department 
                of Agriculture (Parts 2700--2799)
    XXVIII  Office of Operations, Department of Agriculture (Parts 
                2800--2899)
      XXIX  Office of Energy Policy and New Uses, Department of 
                Agriculture (Parts 2900--2999)
       XXX  Office of the Chief Financial Officer, Department of 
                Agriculture (Parts 3000--3099)
      XXXI  Office of Environmental Quality, Department of 
                Agriculture (Parts 3100--3199)
     XXXII  Office of Procurement and Property Management, 
                Department of Agriculture (Parts 3200--3299)
    XXXIII  Office of Transportation, Department of Agriculture 
                (Parts 3300--3399)
     XXXIV  National Institute of Food and Agriculture (Parts 
                3400--3499)
      XXXV  Rural Housing Service, Department of Agriculture 
                (Parts 3500--3599)
     XXXVI  National Agricultural Statistics Service, Department 
                of Agriculture (Parts 3600--3699)
    XXXVII  Economic Research Service, Department of Agriculture 
                (Parts 3700--3799)
   XXXVIII  World Agricultural Outlook Board, Department of 
                Agriculture (Parts 3800--3899)
       XLI  [Reserved]
      XLII  Rural Business-Cooperative Service and Rural Utilities 
                Service, Department of Agriculture (Parts 4200--
                4299)

                    Title 8--Aliens and Nationality

         I  Department of Homeland Security (Immigration and 
                Naturalization) (Parts 1--499)
         V  Executive Office for Immigration Review, Department of 
                Justice (Parts 1000--1399)

[[Page 1066]]

                 Title 9--Animals and Animal Products

         I  Animal and Plant Health Inspection Service, Department 
                of Agriculture (Parts 1--199)
        II  Grain Inspection, Packers and Stockyards 
                Administration (Packers and Stockyards Programs), 
                Department of Agriculture (Parts 200--299)
       III  Food Safety and Inspection Service, Department of 
                Agriculture (Parts 300--599)

                           Title 10--Energy

         I  Nuclear Regulatory Commission (Parts 0--199)
        II  Department of Energy (Parts 200--699)
       III  Department of Energy (Parts 700--999)
         X  Department of Energy (General Provisions) (Parts 
                1000--1099)
      XIII  Nuclear Waste Technical Review Board (Parts 1300--
                1399)
      XVII  Defense Nuclear Facilities Safety Board (Parts 1700--
                1799)
     XVIII  Northeast Interstate Low-Level Radioactive Waste 
                Commission (Parts 1800--1899)

                      Title 11--Federal Elections

         I  Federal Election Commission (Parts 1--9099)
        II  Election Assistance Commission (Parts 9400--9499)

                      Title 12--Banks and Banking

         I  Comptroller of the Currency, Department of the 
                Treasury (Parts 1--199)
        II  Federal Reserve System (Parts 200--299)
       III  Federal Deposit Insurance Corporation (Parts 300--399)
        IV  Export-Import Bank of the United States (Parts 400--
                499)
         V  Office of Thrift Supervision, Department of the 
                Treasury (Parts 500--599)
        VI  Farm Credit Administration (Parts 600--699)
       VII  National Credit Union Administration (Parts 700--799)
      VIII  Federal Financing Bank (Parts 800--899)
        IX  Federal Housing Finance Board (Parts 900--999)
         X  Bureau of Consumer Financial Protection (Parts 1000--
                1099)
        XI  Federal Financial Institutions Examination Council 
                (Parts 1100--1199)
       XII  Federal Housing Finance Agency (Parts 1200--1299)
      XIII  Financial Stability Oversight Council (Parts 1300--
                1399)
       XIV  Farm Credit System Insurance Corporation (Parts 1400--
                1499)
        XV  Department of the Treasury (Parts 1500--1599)
       XVI  Office of Financial Research (Parts 1600--1699)
      XVII  Office of Federal Housing Enterprise Oversight, 
                Department of Housing and Urban Development (Parts 
                1700--1799)

[[Page 1067]]

     XVIII  Community Development Financial Institutions Fund, 
                Department of the Treasury (Parts 1800--1899)

               Title 13--Business Credit and Assistance

         I  Small Business Administration (Parts 1--199)
       III  Economic Development Administration, Department of 
                Commerce (Parts 300--399)
        IV  Emergency Steel Guarantee Loan Board (Parts 400--499)
         V  Emergency Oil and Gas Guaranteed Loan Board (Parts 
                500--599)

                    Title 14--Aeronautics and Space

         I  Federal Aviation Administration, Department of 
                Transportation (Parts 1--199)
        II  Office of the Secretary, Department of Transportation 
                (Aviation Proceedings) (Parts 200--399)
       III  Commercial Space Transportation, Federal Aviation 
                Administration, Department of Transportation 
                (Parts 400--1199)
         V  National Aeronautics and Space Administration (Parts 
                1200--1299)
        VI  Air Transportation System Stabilization (Parts 1300--
                1399)

                 Title 15--Commerce and Foreign Trade

            Subtitle A--Office of the Secretary of Commerce (Parts 
                0--29)
            Subtitle B--Regulations Relating to Commerce and 
                Foreign Trade
         I  Bureau of the Census, Department of Commerce (Parts 
                30--199)
        II  National Institute of Standards and Technology, 
                Department of Commerce (Parts 200--299)
       III  International Trade Administration, Department of 
                Commerce (Parts 300--399)
        IV  Foreign-Trade Zones Board, Department of Commerce 
                (Parts 400--499)
       VII  Bureau of Industry and Security, Department of 
                Commerce (Parts 700--799)
      VIII  Bureau of Economic Analysis, Department of Commerce 
                (Parts 800--899)
        IX  National Oceanic and Atmospheric Administration, 
                Department of Commerce (Parts 900--999)
        XI  Technology Administration, Department of Commerce 
                (Parts 1100--1199)
      XIII  East-West Foreign Trade Board (Parts 1300--1399)
       XIV  Minority Business Development Agency (Parts 1400--
                1499)
            Subtitle C--Regulations Relating to Foreign Trade 
                Agreements

[[Page 1068]]

        XX  Office of the United States Trade Representative 
                (Parts 2000--2099)
            Subtitle D--Regulations Relating to Telecommunications 
                and Information
     XXIII  National Telecommunications and Information 
                Administration, Department of Commerce (Parts 
                2300--2399)

                    Title 16--Commercial Practices

         I  Federal Trade Commission (Parts 0--999)
        II  Consumer Product Safety Commission (Parts 1000--1799)

             Title 17--Commodity and Securities Exchanges

         I  Commodity Futures Trading Commission (Parts 1--199)
        II  Securities and Exchange Commission (Parts 200--399)
        IV  Department of the Treasury (Parts 400--499)

          Title 18--Conservation of Power and Water Resources

         I  Federal Energy Regulatory Commission, Department of 
                Energy (Parts 1--399)
       III  Delaware River Basin Commission (Parts 400--499)
        VI  Water Resources Council (Parts 700--799)
      VIII  Susquehanna River Basin Commission (Parts 800--899)
      XIII  Tennessee Valley Authority (Parts 1300--1399)

                       Title 19--Customs Duties

         I  U.S. Customs and Border Protection, Department of 
                Homeland Security; Department of the Treasury 
                (Parts 0--199)
        II  United States International Trade Commission (Parts 
                200--299)
       III  International Trade Administration, Department of 
                Commerce (Parts 300--399)
        IV  U.S. Immigration and Customs Enforcement, Department 
                of Homeland Security (Parts 400--599)

                     Title 20--Employees' Benefits

         I  Office of Workers' Compensation Programs, Department 
                of Labor (Parts 1--199)
        II  Railroad Retirement Board (Parts 200--399)
       III  Social Security Administration (Parts 400--499)
        IV  Employees' Compensation Appeals Board, Department of 
                Labor (Parts 500--599)
         V  Employment and Training Administration, Department of 
                Labor (Parts 600--699)

[[Page 1069]]

        VI  Office of Workers' Compensation Programs, Department 
                of Labor (Parts 700--799)
       VII  Benefits Review Board, Department of Labor (Parts 
                800--899)
      VIII  Joint Board for the Enrollment of Actuaries (Parts 
                900--999)
        IX  Office of the Assistant Secretary for Veterans' 
                Employment and Training Service, Department of 
                Labor (Parts 1000--1099)

                       Title 21--Food and Drugs

         I  Food and Drug Administration, Department of Health and 
                Human Services (Parts 1--1299)
        II  Drug Enforcement Administration, Department of Justice 
                (Parts 1300--1399)
       III  Office of National Drug Control Policy (Parts 1400--
                1499)

                      Title 22--Foreign Relations

         I  Department of State (Parts 1--199)
        II  Agency for International Development (Parts 200--299)
       III  Peace Corps (Parts 300--399)
        IV  International Joint Commission, United States and 
                Canada (Parts 400--499)
         V  Broadcasting Board of Governors (Parts 500--599)
       VII  Overseas Private Investment Corporation (Parts 700--
                799)
        IX  Foreign Service Grievance Board (Parts 900--999)
         X  Inter-American Foundation (Parts 1000--1099)
        XI  International Boundary and Water Commission, United 
                States and Mexico, United States Section (Parts 
                1100--1199)
       XII  United States International Development Cooperation 
                Agency (Parts 1200--1299)
      XIII  Millennium Challenge Corporation (Parts 1300--1399)
       XIV  Foreign Service Labor Relations Board; Federal Labor 
                Relations Authority; General Counsel of the 
                Federal Labor Relations Authority; and the Foreign 
                Service Impasse Disputes Panel (Parts 1400--1499)
        XV  African Development Foundation (Parts 1500--1599)
       XVI  Japan-United States Friendship Commission (Parts 
                1600--1699)
      XVII  United States Institute of Peace (Parts 1700--1799)

                          Title 23--Highways

         I  Federal Highway Administration, Department of 
                Transportation (Parts 1--999)
        II  National Highway Traffic Safety Administration and 
                Federal Highway Administration, Department of 
                Transportation (Parts 1200--1299)
       III  National Highway Traffic Safety Administration, 
                Department of Transportation (Parts 1300--1399)

[[Page 1070]]

                Title 24--Housing and Urban Development

            Subtitle A--Office of the Secretary, Department of 
                Housing and Urban Development (Parts 0--99)
            Subtitle B--Regulations Relating to Housing and Urban 
                Development
         I  Office of Assistant Secretary for Equal Opportunity, 
                Department of Housing and Urban Development (Parts 
                100--199)
        II  Office of Assistant Secretary for Housing-Federal 
                Housing Commissioner, Department of Housing and 
                Urban Development (Parts 200--299)
       III  Government National Mortgage Association, Department 
                of Housing and Urban Development (Parts 300--399)
        IV  Office of Housing and Office of Multifamily Housing 
                Assistance Restructuring, Department of Housing 
                and Urban Development (Parts 400--499)
         V  Office of Assistant Secretary for Community Planning 
                and Development, Department of Housing and Urban 
                Development (Parts 500--599)
        VI  Office of Assistant Secretary for Community Planning 
                and Development, Department of Housing and Urban 
                Development (Parts 600--699) [Reserved]
       VII  Office of the Secretary, Department of Housing and 
                Urban Development (Housing Assistance Programs and 
                Public and Indian Housing Programs) (Parts 700--
                799)
      VIII  Office of the Assistant Secretary for Housing--Federal 
                Housing Commissioner, Department of Housing and 
                Urban Development (Section 8 Housing Assistance 
                Programs, Section 202 Direct Loan Program, Section 
                202 Supportive Housing for the Elderly Program and 
                Section 811 Supportive Housing for Persons With 
                Disabilities Program) (Parts 800--899)
        IX  Office of Assistant Secretary for Public and Indian 
                Housing, Department of Housing and Urban 
                Development (Parts 900--1699)
         X  Office of Assistant Secretary for Housing--Federal 
                Housing Commissioner, Department of Housing and 
                Urban Development (Interstate Land Sales 
                Registration Program) (Parts 1700--1799)
       XII  Office of Inspector General, Department of Housing and 
                Urban Development (Parts 2000--2099)
        XV  Emergency Mortgage Insurance and Loan Programs, 
                Department of Housing and Urban Development (Parts 
                2700--2799) [Reserved]
        XX  Office of Assistant Secretary for Housing--Federal 
                Housing Commissioner, Department of Housing and 
                Urban Development (Parts 3200--3899)
      XXIV  Board of Directors of the HOPE for Homeowners Program 
                (Parts 4000--4099) [Reserved]
       XXV  Neighborhood Reinvestment Corporation (Parts 4100--
                4199)

                           Title 25--Indians

         I  Bureau of Indian Affairs, Department of the Interior 
                (Parts 1--299)

[[Page 1071]]

        II  Indian Arts and Crafts Board, Department of the 
                Interior (Parts 300--399)
       III  National Indian Gaming Commission, Department of the 
                Interior (Parts 500--599)
        IV  Office of Navajo and Hopi Indian Relocation (Parts 
                700--799)
         V  Bureau of Indian Affairs, Department of the Interior, 
                and Indian Health Service, Department of Health 
                and Human Services (Part 900)
        VI  Office of the Assistant Secretary-Indian Affairs, 
                Department of the Interior (Parts 1000--1199)
       VII  Office of the Special Trustee for American Indians, 
                Department of the Interior (Parts 1200--1299)

                      Title 26--Internal Revenue

         I  Internal Revenue Service, Department of the Treasury 
                (Parts 1--End)

           Title 27--Alcohol, Tobacco Products and Firearms

         I  Alcohol and Tobacco Tax and Trade Bureau, Department 
                of the Treasury (Parts 1--399)
        II  Bureau of Alcohol, Tobacco, Firearms, and Explosives, 
                Department of Justice (Parts 400--699)

                   Title 28--Judicial Administration

         I  Department of Justice (Parts 0--299)
       III  Federal Prison Industries, Inc., Department of Justice 
                (Parts 300--399)
         V  Bureau of Prisons, Department of Justice (Parts 500--
                599)
        VI  Offices of Independent Counsel, Department of Justice 
                (Parts 600--699)
       VII  Office of Independent Counsel (Parts 700--799)
      VIII  Court Services and Offender Supervision Agency for the 
                District of Columbia (Parts 800--899)
        IX  National Crime Prevention and Privacy Compact Council 
                (Parts 900--999)
        XI  Department of Justice and Department of State (Parts 
                1100--1199)

                            Title 29--Labor

            Subtitle A--Office of the Secretary of Labor (Parts 
                0--99)
            Subtitle B--Regulations Relating to Labor
         I  National Labor Relations Board (Parts 100--199)
        II  Office of Labor-Management Standards, Department of 
                Labor (Parts 200--299)
       III  National Railroad Adjustment Board (Parts 300--399)

[[Page 1072]]

        IV  Office of Labor-Management Standards, Department of 
                Labor (Parts 400--499)
         V  Wage and Hour Division, Department of Labor (Parts 
                500--899)
        IX  Construction Industry Collective Bargaining Commission 
                (Parts 900--999)
         X  National Mediation Board (Parts 1200--1299)
       XII  Federal Mediation and Conciliation Service (Parts 
                1400--1499)
       XIV  Equal Employment Opportunity Commission (Parts 1600--
                1699)
      XVII  Occupational Safety and Health Administration, 
                Department of Labor (Parts 1900--1999)
        XX  Occupational Safety and Health Review Commission 
                (Parts 2200--2499)
       XXV  Employee Benefits Security Administration, Department 
                of Labor (Parts 2500--2599)
     XXVII  Federal Mine Safety and Health Review Commission 
                (Parts 2700--2799)
        XL  Pension Benefit Guaranty Corporation (Parts 4000--
                4999)

                      Title 30--Mineral Resources

         I  Mine Safety and Health Administration, Department of 
                Labor (Parts 1--199)
        II  Bureau of Safety and Environmental Enforcement, 
                Department of the Interior (Parts 200--299)
        IV  Geological Survey, Department of the Interior (Parts 
                400--499)
         V  Bureau of Ocean Energy Management, Department of the 
                Interior (Parts 500--599)
       VII  Office of Surface Mining Reclamation and Enforcement, 
                Department of the Interior (Parts 700--999)
       XII  Office of Natural Resources Revenue, Department of the 
                Interior (Parts 1200--1299)

                 Title 31--Money and Finance: Treasury

            Subtitle A--Office of the Secretary of the Treasury 
                (Parts 0--50)
            Subtitle B--Regulations Relating to Money and Finance
         I  Monetary Offices, Department of the Treasury (Parts 
                51--199)
        II  Fiscal Service, Department of the Treasury (Parts 
                200--399)
        IV  Secret Service, Department of the Treasury (Parts 
                400--499)
         V  Office of Foreign Assets Control, Department of the 
                Treasury (Parts 500--599)
        VI  Bureau of Engraving and Printing, Department of the 
                Treasury (Parts 600--699)
       VII  Federal Law Enforcement Training Center, Department of 
                the Treasury (Parts 700--799)
      VIII  Office of International Investment, Department of the 
                Treasury (Parts 800--899)

[[Page 1073]]

        IX  Federal Claims Collection Standards (Department of the 
                Treasury--Department of Justice) (Parts 900--999)
         X  Financial Crimes Enforcement Network, Department of 
                the Treasury (Parts 1000--1099)

                      Title 32--National Defense

            Subtitle A--Department of Defense
         I  Office of the Secretary of Defense (Parts 1--399)
         V  Department of the Army (Parts 400--699)
        VI  Department of the Navy (Parts 700--799)
       VII  Department of the Air Force (Parts 800--1099)
            Subtitle B--Other Regulations Relating to National 
                Defense
       XII  Defense Logistics Agency (Parts 1200--1299)
       XVI  Selective Service System (Parts 1600--1699)
      XVII  Office of the Director of National Intelligence (Parts 
                1700--1799)
     XVIII  National Counterintelligence Center (Parts 1800--1899)
       XIX  Central Intelligence Agency (Parts 1900--1999)
        XX  Information Security Oversight Office, National 
                Archives and Records Administration (Parts 2000--
                2099)
       XXI  National Security Council (Parts 2100--2199)
      XXIV  Office of Science and Technology Policy (Parts 2400--
                2499)
     XXVII  Office for Micronesian Status Negotiations (Parts 
                2700--2799)
    XXVIII  Office of the Vice President of the United States 
                (Parts 2800--2899)

               Title 33--Navigation and Navigable Waters

         I  Coast Guard, Department of Homeland Security (Parts 
                1--199)
        II  Corps of Engineers, Department of the Army (Parts 
                200--399)
        IV  Saint Lawrence Seaway Development Corporation, 
                Department of Transportation (Parts 400--499)

                          Title 34--Education

            Subtitle A--Office of the Secretary, Department of 
                Education (Parts 1--99)
            Subtitle B--Regulations of the Offices of the 
                Department of Education
         I  Office for Civil Rights, Department of Education 
                (Parts 100--199)
        II  Office of Elementary and Secondary Education, 
                Department of Education (Parts 200--299)
       III  Office of Special Education and Rehabilitative 
                Services, Department of Education (Parts 300--399)
        IV  Office of Career, Technical, and Adult Education, 
                Department of Education (Parts 400--499)

[[Page 1074]]

         V  Office of Bilingual Education and Minority Languages 
                Affairs, Department of Education (Parts 500--599)
        VI  Office of Postsecondary Education, Department of 
                Education (Parts 600--699)
       VII  Office of Educational Research and Improvement, 
                Department of Education (Parts 700--799)[Reserved]
            Subtitle C--Regulations Relating to Education
        XI  National Institute for Literacy [Reserved]
       XII  National Council on Disability (Parts 1200--1299)

                          Title 35 [Reserved]

             Title 36--Parks, Forests, and Public Property

         I  National Park Service, Department of the Interior 
                (Parts 1--199)
        II  Forest Service, Department of Agriculture (Parts 200--
                299)
       III  Corps of Engineers, Department of the Army (Parts 
                300--399)
        IV  American Battle Monuments Commission (Parts 400--499)
         V  Smithsonian Institution (Parts 500--599)
        VI  [Reserved]
       VII  Library of Congress (Parts 700--799)
      VIII  Advisory Council on Historic Preservation (Parts 800--
                899)
        IX  Pennsylvania Avenue Development Corporation (Parts 
                900--999)
         X  Presidio Trust (Parts 1000--1099)
        XI  Architectural and Transportation Barriers Compliance 
                Board (Parts 1100--1199)
       XII  National Archives and Records Administration (Parts 
                1200--1299)
        XV  Oklahoma City National Memorial Trust (Parts 1500--
                1599)
       XVI  Morris K. Udall Scholarship and Excellence in National 
                Environmental Policy Foundation (Parts 1600--1699)

             Title 37--Patents, Trademarks, and Copyrights

         I  United States Patent and Trademark Office, Department 
                of Commerce (Parts 1--199)
        II  U.S. Copyright Office, Library of Congress (Parts 
                200--299)
       III  Copyright Royalty Board, Library of Congress (Parts 
                300--399)
        IV  Assistant Secretary for Technology Policy, Department 
                of Commerce (Parts 400--599)

           Title 38--Pensions, Bonuses, and Veterans' Relief

         I  Department of Veterans Affairs (Parts 0--199)
        II  Armed Forces Retirement Home (Parts 200--299)

[[Page 1075]]

                       Title 39--Postal Service

         I  United States Postal Service (Parts 1--999)
       III  Postal Regulatory Commission (Parts 3000--3099)

                  Title 40--Protection of Environment

         I  Environmental Protection Agency (Parts 1--1099)
        IV  Environmental Protection Agency and Department of 
                Justice (Parts 1400--1499)
         V  Council on Environmental Quality (Parts 1500--1599)
        VI  Chemical Safety and Hazard Investigation Board (Parts 
                1600--1699)
       VII  Environmental Protection Agency and Department of 
                Defense; Uniform National Discharge Standards for 
                Vessels of the Armed Forces (Parts 1700--1799)
      VIII  Gulf Coast Ecosystem Restoration Council (Parts 1800--
                1899)

          Title 41--Public Contracts and Property Management

            Subtitle A--Federal Procurement Regulations System 
                [Note]
            Subtitle B--Other Provisions Relating to Public 
                Contracts
        50  Public Contracts, Department of Labor (Parts 50-1--50-
                999)
        51  Committee for Purchase From People Who Are Blind or 
                Severely Disabled (Parts 51-1--51-99)
        60  Office of Federal Contract Compliance Programs, Equal 
                Employment Opportunity, Department of Labor (Parts 
                60-1--60-999)
        61  Office of the Assistant Secretary for Veterans' 
                Employment and Training Service, Department of 
                Labor (Parts 61-1--61-999)
   62--100  [Reserved]
            Subtitle C--Federal Property Management Regulations 
                System
       101  Federal Property Management Regulations (Parts 101-1--
                101-99)
       102  Federal Management Regulation (Parts 102-1--102-299)
  103--104  [Reserved]
       105  General Services Administration (Parts 105-1--105-999)
       109  Department of Energy Property Management Regulations 
                (Parts 109-1--109-99)
       114  Department of the Interior (Parts 114-1--114-99)
       115  Environmental Protection Agency (Parts 115-1--115-99)
       128  Department of Justice (Parts 128-1--128-99)
  129--200  [Reserved]
            Subtitle D--Other Provisions Relating to Property 
                Management [Reserved]
            Subtitle E--Federal Information Resources Management 
                Regulations System [Reserved]
            Subtitle F--Federal Travel Regulation System
       300  General (Parts 300-1--300-99)
       301  Temporary Duty (TDY) Travel Allowances (Parts 301-1--
                301-99)

[[Page 1076]]

       302  Relocation Allowances (Parts 302-1--302-99)
       303  Payment of Expenses Connected with the Death of 
                Certain Employees (Part 303-1--303-99)
       304  Payment of Travel Expenses from a Non-Federal Source 
                (Parts 304-1--304-99)

                        Title 42--Public Health

         I  Public Health Service, Department of Health and Human 
                Services (Parts 1--199)
        IV  Centers for Medicare & Medicaid Services, Department 
                of Health and Human Services (Parts 400--599)
         V  Office of Inspector General-Health Care, Department of 
                Health and Human Services (Parts 1000--1999)

                   Title 43--Public Lands: Interior

            Subtitle A--Office of the Secretary of the Interior 
                (Parts 1--199)
            Subtitle B--Regulations Relating to Public Lands
         I  Bureau of Reclamation, Department of the Interior 
                (Parts 400--999)
        II  Bureau of Land Management, Department of the Interior 
                (Parts 1000--9999)
       III  Utah Reclamation Mitigation and Conservation 
                Commission (Parts 10000--10099)

             Title 44--Emergency Management and Assistance

         I  Federal Emergency Management Agency, Department of 
                Homeland Security (Parts 0--399)
        IV  Department of Commerce and Department of 
                Transportation (Parts 400--499)

                       Title 45--Public Welfare

            Subtitle A--Department of Health and Human Services 
                (Parts 1--199)
            Subtitle B--Regulations Relating to Public Welfare
        II  Office of Family Assistance (Assistance Programs), 
                Administration for Children and Families, 
                Department of Health and Human Services (Parts 
                200--299)
       III  Office of Child Support Enforcement (Child Support 
                Enforcement Program), Administration for Children 
                and Families, Department of Health and Human 
                Services (Parts 300--399)
        IV  Office of Refugee Resettlement, Administration for 
                Children and Families, Department of Health and 
                Human Services (Parts 400--499)
         V  Foreign Claims Settlement Commission of the United 
                States, Department of Justice (Parts 500--599)

[[Page 1077]]

        VI  National Science Foundation (Parts 600--699)
       VII  Commission on Civil Rights (Parts 700--799)
      VIII  Office of Personnel Management (Parts 800--899)
         X  Office of Community Services, Administration for 
                Children and Families, Department of Health and 
                Human Services (Parts 1000--1099)
        XI  National Foundation on the Arts and the Humanities 
                (Parts 1100--1199)
       XII  Corporation for National and Community Service (Parts 
                1200--1299)
      XIII  Office of Human Development Services, Department of 
                Health and Human Services (Parts 1300--1399)
       XVI  Legal Services Corporation (Parts 1600--1699)
      XVII  National Commission on Libraries and Information 
                Science (Parts 1700--1799)
     XVIII  Harry S. Truman Scholarship Foundation (Parts 1800--
                1899)
       XXI  Commission on Fine Arts (Parts 2100--2199)
     XXIII  Arctic Research Commission (Part 2301)
      XXIV  James Madison Memorial Fellowship Foundation (Parts 
                2400--2499)
       XXV  Corporation for National and Community Service (Parts 
                2500--2599)

                          Title 46--Shipping

         I  Coast Guard, Department of Homeland Security (Parts 
                1--199)
        II  Maritime Administration, Department of Transportation 
                (Parts 200--399)
       III  Coast Guard (Great Lakes Pilotage), Department of 
                Homeland Security (Parts 400--499)
        IV  Federal Maritime Commission (Parts 500--599)

                      Title 47--Telecommunication

         I  Federal Communications Commission (Parts 0--199)
        II  Office of Science and Technology Policy and National 
                Security Council (Parts 200--299)
       III  National Telecommunications and Information 
                Administration, Department of Commerce (Parts 
                300--399)
        IV  National Telecommunications and Information 
                Administration, Department of Commerce, and 
                National Highway Traffic Safety Administration, 
                Department of Transportation (Parts 400--499)

           Title 48--Federal Acquisition Regulations System

         1  Federal Acquisition Regulation (Parts 1--99)
         2  Defense Acquisition Regulations System, Department of 
                Defense (Parts 200--299)

[[Page 1078]]

         3  Health and Human Services (Parts 300--399)
         4  Department of Agriculture (Parts 400--499)
         5  General Services Administration (Parts 500--599)
         6  Department of State (Parts 600--699)
         7  Agency for International Development (Parts 700--799)
         8  Department of Veterans Affairs (Parts 800--899)
         9  Department of Energy (Parts 900--999)
        10  Department of the Treasury (Parts 1000--1099)
        12  Department of Transportation (Parts 1200--1299)
        13  Department of Commerce (Parts 1300--1399)
        14  Department of the Interior (Parts 1400--1499)
        15  Environmental Protection Agency (Parts 1500--1599)
        16  Office of Personnel Management, Federal Employees 
                Health Benefits Acquisition Regulation (Parts 
                1600--1699)
        17  Office of Personnel Management (Parts 1700--1799)
        18  National Aeronautics and Space Administration (Parts 
                1800--1899)
        19  Broadcasting Board of Governors (Parts 1900--1999)
        20  Nuclear Regulatory Commission (Parts 2000--2099)
        21  Office of Personnel Management, Federal Employees 
                Group Life Insurance Federal Acquisition 
                Regulation (Parts 2100--2199)
        23  Social Security Administration (Parts 2300--2399)
        24  Department of Housing and Urban Development (Parts 
                2400--2499)
        25  National Science Foundation (Parts 2500--2599)
        28  Department of Justice (Parts 2800--2899)
        29  Department of Labor (Parts 2900--2999)
        30  Department of Homeland Security, Homeland Security 
                Acquisition Regulation (HSAR) (Parts 3000--3099)
        34  Department of Education Acquisition Regulation (Parts 
                3400--3499)
        51  Department of the Army Acquisition Regulations (Parts 
                5100--5199)
        52  Department of the Navy Acquisition Regulations (Parts 
                5200--5299)
        53  Department of the Air Force Federal Acquisition 
                Regulation Supplement (Parts 5300--5399) 
                [Reserved]
        54  Defense Logistics Agency, Department of Defense (Parts 
                5400--5499)
        57  African Development Foundation (Parts 5700--5799)
        61  Civilian Board of Contract Appeals, General Services 
                Administration (Parts 6100--6199)
        63  Department of Transportation Board of Contract Appeals 
                (Parts 6300--6399)
        99  Cost Accounting Standards Board, Office of Federal 
                Procurement Policy, Office of Management and 
                Budget (Parts 9900--9999)

[[Page 1079]]

                       Title 49--Transportation

            Subtitle A--Office of the Secretary of Transportation 
                (Parts 1--99)
            Subtitle B--Other Regulations Relating to 
                Transportation
         I  Pipeline and Hazardous Materials Safety 
                Administration, Department of Transportation 
                (Parts 100--199)
        II  Federal Railroad Administration, Department of 
                Transportation (Parts 200--299)
       III  Federal Motor Carrier Safety Administration, 
                Department of Transportation (Parts 300--399)
        IV  Coast Guard, Department of Homeland Security (Parts 
                400--499)
         V  National Highway Traffic Safety Administration, 
                Department of Transportation (Parts 500--599)
        VI  Federal Transit Administration, Department of 
                Transportation (Parts 600--699)
       VII  National Railroad Passenger Corporation (AMTRAK) 
                (Parts 700--799)
      VIII  National Transportation Safety Board (Parts 800--999)
         X  Surface Transportation Board, Department of 
                Transportation (Parts 1000--1399)
        XI  Research and Innovative Technology Administration, 
                Department of Transportation (Parts 1400--1499) 
                [Reserved]
       XII  Transportation Security Administration, Department of 
                Homeland Security (Parts 1500--1699)

                   Title 50--Wildlife and Fisheries

         I  United States Fish and Wildlife Service, Department of 
                the Interior (Parts 1--199)
        II  National Marine Fisheries Service, National Oceanic 
                and Atmospheric Administration, Department of 
                Commerce (Parts 200--299)
       III  International Fishing and Related Activities (Parts 
                300--399)
        IV  Joint Regulations (United States Fish and Wildlife 
                Service, Department of the Interior and National 
                Marine Fisheries Service, National Oceanic and 
                Atmospheric Administration, Department of 
                Commerce); Endangered Species Committee 
                Regulations (Parts 400--499)
         V  Marine Mammal Commission (Parts 500--599)
        VI  Fishery Conservation and Management, National Oceanic 
                and Atmospheric Administration, Department of 
                Commerce (Parts 600--699)

[[Page 1081]]





           Alphabetical List of Agencies Appearing in the CFR




                     (Revised as of January 1, 2015)

                                                  CFR Title, Subtitle or 
                     Agency                               Chapter

Administrative Committee of the Federal Register  1, I
Administrative Conference of the United States    1, III
Advisory Council on Historic Preservation         36, VIII
Advocacy and Outreach, Office of                  7, XXV
Afghanistan Reconstruction, Special Inspector     22, LXXXIII
     General for
African Development Foundation                    22, XV
  Federal Acquisition Regulation                  48, 57
Agency for International Development              2, VII; 22, II
  Federal Acquisition Regulation                  48, 7
Agricultural Marketing Service                    7, I, IX, X, XI
Agricultural Research Service                     7, V
Agriculture Department                            2, IV; 5, LXXIII
  Advocacy and Outreach, Office of                7, XXV
  Agricultural Marketing Service                  7, I, IX, X, XI
  Agricultural Research Service                   7, V
  Animal and Plant Health Inspection Service      7, III; 9, I
  Chief Financial Officer, Office of              7, XXX
  Commodity Credit Corporation                    7, XIV
  Economic Research Service                       7, XXXVII
  Energy Policy and New Uses, Office of           2, IX; 7, XXIX
  Environmental Quality, Office of                7, XXXI
  Farm Service Agency                             7, VII, XVIII
  Federal Acquisition Regulation                  48, 4
  Federal Crop Insurance Corporation              7, IV
  Food and Nutrition Service                      7, II
  Food Safety and Inspection Service              9, III
  Foreign Agricultural Service                    7, XV
  Forest Service                                  36, II
  Grain Inspection, Packers and Stockyards        7, VIII; 9, II
       Administration
  Information Resources Management, Office of     7, XXVII
  Inspector General, Office of                    7, XXVI
  National Agricultural Library                   7, XLI
  National Agricultural Statistics Service        7, XXXVI
  National Institute of Food and Agriculture      7, XXXIV
  Natural Resources Conservation Service          7, VI
  Operations, Office of                           7, XXVIII
  Procurement and Property Management, Office of  7, XXXII
  Rural Business-Cooperative Service              7, XVIII, XLII, L
  Rural Development Administration                7, XLII
  Rural Housing Service                           7, XVIII, XXXV, L
  Rural Telephone Bank                            7, XVI
  Rural Utilities Service                         7, XVII, XVIII, XLII, L
  Secretary of Agriculture, Office of             7, Subtitle A
  Transportation, Office of                       7, XXXIII
  World Agricultural Outlook Board                7, XXXVIII
Air Force Department                              32, VII
  Federal Acquisition Regulation Supplement       48, 53
Air Transportation Stabilization Board            14, VI
Alcohol and Tobacco Tax and Trade Bureau          27, I
Alcohol, Tobacco, Firearms, and Explosives,       27, II
     Bureau of
AMTRAK                                            49, VII
American Battle Monuments Commission              36, IV
American Indians, Office of the Special Trustee   25, VII

[[Page 1082]]

Animal and Plant Health Inspection Service        7, III; 9, I
Appalachian Regional Commission                   5, IX
Architectural and Transportation Barriers         36, XI
     Compliance Board
Arctic Research Commission                        45, XXIII
Armed Forces Retirement Home                      5, XI
Army Department                                   32, V
  Engineers, Corps of                             33, II; 36, III
  Federal Acquisition Regulation                  48, 51
Bilingual Education and Minority Languages        34, V
     Affairs, Office of
Blind or Severely Disabled, Committee for         41, 51
     Purchase from People Who Are
Broadcasting Board of Governors                   22, V
  Federal Acquisition Regulation                  48, 19
Bureau of Ocean Energy Management, Regulation,    30, II
     and Enforcement
Census Bureau                                     15, I
Centers for Medicare & Medicaid Services          42, IV
Central Intelligence Agency                       32, XIX
Chemical Safety and Hazardous Investigation       40, VI
     Board
Chief Financial Officer, Office of                7, XXX
Child Support Enforcement, Office of              45, III
Children and Families, Administration for         45, II, III, IV, X
Civil Rights, Commission on                       5, LXVIII; 45, VII
Civil Rights, Office for                          34, I
Council of the Inspectors General on Integrity    5, XCVIII
     and Efficiency
Court Services and Offender Supervision Agency    5, LXX
     for the District of Columbia
Coast Guard                                       33, I; 46, I; 49, IV
Coast Guard (Great Lakes Pilotage)                46, III
Commerce Department                               2, XIII; 44, IV; 50, VI
  Census Bureau                                   15, I
  Economic Analysis, Bureau of                    15, VIII
  Economic Development Administration             13, III
  Emergency Management and Assistance             44, IV
  Federal Acquisition Regulation                  48, 13
  Foreign-Trade Zones Board                       15, IV
  Industry and Security, Bureau of                15, VII
  International Trade Administration              15, III; 19, III
  National Institute of Standards and Technology  15, II
  National Marine Fisheries Service               50, II, IV
  National Oceanic and Atmospheric                15, IX; 50, II, III, IV, 
       Administration                             VI
  National Telecommunications and Information     15, XXIII; 47, III, IV
       Administration
  National Weather Service                        15, IX
  Patent and Trademark Office, United States      37, I
  Productivity, Technology and Innovation,        37, IV
       Assistant Secretary for
  Secretary of Commerce, Office of                15, Subtitle A
  Technology Administration                       15, XI
  Technology Policy, Assistant Secretary for      37, IV
Commercial Space Transportation                   14, III
Commodity Credit Corporation                      7, XIV
Commodity Futures Trading Commission              5, XLI; 17, I
Community Planning and Development, Office of     24, V, VI
     Assistant Secretary for
Community Services, Office of                     45, X
Comptroller of the Currency                       12, I
Construction Industry Collective Bargaining       29, IX
     Commission
Consumer Financial Protection Bureau              5, LXXXIV; 12, X
Consumer Product Safety Commission                5, LXXI; 16, II
Copyright Royalty Board                           37, III
Corporation for National and Community Service    2, XXII; 45, XII, XXV
Cost Accounting Standards Board                   48, 99
Council on Environmental Quality                  40, V
Court Services and Offender Supervision Agency    5, LXX; 28, VIII
     for the District of Columbia
Customs and Border Protection                     19, I

[[Page 1083]]

Defense Contract Audit Agency                     32, I
Defense Department                                2, XI; 5, XXVI; 32, 
                                                  Subtitle A; 40, VII
  Advanced Research Projects Agency               32, I
  Air Force Department                            32, VII
  Army Department                                 32, V; 33, II; 36, III, 
                                                  48, 51
  Defense Acquisition Regulations System          48, 2
  Defense Intelligence Agency                     32, I
  Defense Logistics Agency                        32, I, XII; 48, 54
  Engineers, Corps of                             33, II; 36, III
  National Imagery and Mapping Agency             32, I
  Navy Department                                 32, VI; 48, 52
  Secretary of Defense, Office of                 2, XI; 32, I
Defense Contract Audit Agency                     32, I
Defense Intelligence Agency                       32, I
Defense Logistics Agency                          32, XII; 48, 54
Defense Nuclear Facilities Safety Board           10, XVII
Delaware River Basin Commission                   18, III
District of Columbia, Court Services and          5, LXX; 28, VIII
     Offender Supervision Agency for the
Drug Enforcement Administration                   21, II
East-West Foreign Trade Board                     15, XIII
Economic Analysis, Bureau of                      15, VIII
Economic Development Administration               13, III
Economic Research Service                         7, XXXVII
Education, Department of                          2, XXXIV; 5, LIII
  Bilingual Education and Minority Languages      34, V
       Affairs, Office of
  Civil Rights, Office for                        34, I
  Educational Research and Improvement, Office    34, VII
       of
  Elementary and Secondary Education, Office of   34, II
  Federal Acquisition Regulation                  48, 34
  Postsecondary Education, Office of              34, VI
  Secretary of Education, Office of               34, Subtitle A
  Special Education and Rehabilitative Services,  34, III
       Office of
  Vocational and Adult Education, Office of       34, IV
Educational Research and Improvement, Office of   34, VII
Election Assistance Commission                    2, LVIII; 11, II
Elementary and Secondary Education, Office of     34, II
Emergency Oil and Gas Guaranteed Loan Board       13, V
Emergency Steel Guarantee Loan Board              13, IV
Employee Benefits Security Administration         29, XXV
Employees' Compensation Appeals Board             20, IV
Employees Loyalty Board                           5, V
Employment and Training Administration            20, V
Employment Standards Administration               20, VI
Endangered Species Committee                      50, IV
Energy, Department of                             2, IX; 5, XXIII; 10, II, 
                                                  III, X
  Federal Acquisition Regulation                  48, 9
  Federal Energy Regulatory Commission            5, XXIV; 18, I
  Property Management Regulations                 41, 109
Energy, Office of                                 7, XXIX
Engineers, Corps of                               33, II; 36, III
Engraving and Printing, Bureau of                 31, VI
Environmental Protection Agency                   2, XV; 5, LIV; 40, I, IV, 
                                                  VII
  Federal Acquisition Regulation                  48, 15
  Property Management Regulations                 41, 115
Environmental Quality, Office of                  7, XXXI
Equal Employment Opportunity Commission           5, LXII; 29, XIV
Equal Opportunity, Office of Assistant Secretary  24, I
     for
Executive Office of the President                 3, I
  Administration, Office of                       5, XV
  Environmental Quality, Council on               40, V
  Management and Budget, Office of                2, Subtitle A; 5, III, 
                                                  LXXVII; 14, VI; 48, 99

[[Page 1084]]

  National Drug Control Policy, Office of         21, III
  National Security Council                       32, XXI; 47, 2
  Presidential Documents                          3
  Science and Technology Policy, Office of        32, XXIV; 47, II
  Trade Representative, Office of the United      15, XX
       States
Export-Import Bank of the United States           2, XXXV; 5, LII; 12, IV
Family Assistance, Office of                      45, II
Farm Credit Administration                        5, XXXI; 12, VI
Farm Credit System Insurance Corporation          5, XXX; 12, XIV
Farm Service Agency                               7, VII, XVIII
Federal Acquisition Regulation                    48, 1
Federal Aviation Administration                   14, I
  Commercial Space Transportation                 14, III
Federal Claims Collection Standards               31, IX
Federal Communications Commission                 5, XXIX; 47, I
Federal Contract Compliance Programs, Office of   41, 60
Federal Crop Insurance Corporation                7, IV
Federal Deposit Insurance Corporation             5, XXII; 12, III
Federal Election Commission                       5, XXXVII; 11, I
Federal Emergency Management Agency               44, I
Federal Employees Group Life Insurance Federal    48, 21
     Acquisition Regulation
Federal Employees Health Benefits Acquisition     48, 16
     Regulation
Federal Energy Regulatory Commission              5, XXIV; 18, I
Federal Financial Institutions Examination        12, XI
     Council
Federal Financing Bank                            12, VIII
Federal Highway Administration                    23, I, II
Federal Home Loan Mortgage Corporation            1, IV
Federal Housing Enterprise Oversight Office       12, XVII
Federal Housing Finance Agency                    5, LXXX; 12, XII
Federal Housing Finance Board                     12, IX
Federal Labor Relations Authority                 5, XIV, XLIX; 22, XIV
Federal Law Enforcement Training Center           31, VII
Federal Management Regulation                     41, 102
Federal Maritime Commission                       46, IV
Federal Mediation and Conciliation Service        29, XII
Federal Mine Safety and Health Review Commission  5, LXXIV; 29, XXVII
Federal Motor Carrier Safety Administration       49, III
Federal Prison Industries, Inc.                   28, III
Federal Procurement Policy Office                 48, 99
Federal Property Management Regulations           41, 101
Federal Railroad Administration                   49, II
Federal Register, Administrative Committee of     1, I
Federal Register, Office of                       1, II
Federal Reserve System                            12, II
  Board of Governors                              5, LVIII
Federal Retirement Thrift Investment Board        5, VI, LXXVI
Federal Service Impasses Panel                    5, XIV
Federal Trade Commission                          5, XLVII; 16, I
Federal Transit Administration                    49, VI
Federal Travel Regulation System                  41, Subtitle F
Financial Crimes Enforcement Network              31, X
Financial Research Office                         12, XVI
Financial Stability Oversight Council             12, XIII
Fine Arts, Commission on                          45, XXI
Fiscal Service                                    31, II
Fish and Wildlife Service, United States          50, I, IV
Food and Drug Administration                      21, I
Food and Nutrition Service                        7, II
Food Safety and Inspection Service                9, III
Foreign Agricultural Service                      7, XV
Foreign Assets Control, Office of                 31, V
Foreign Claims Settlement Commission of the       45, V
     United States
Foreign Service Grievance Board                   22, IX
Foreign Service Impasse Disputes Panel            22, XIV
Foreign Service Labor Relations Board             22, XIV
Foreign-Trade Zones Board                         15, IV

[[Page 1085]]

Forest Service                                    36, II
General Services Administration                   5, LVII; 41, 105
  Contract Appeals, Board of                      48, 61
  Federal Acquisition Regulation                  48, 5
  Federal Management Regulation                   41, 102
  Federal Property Management Regulations         41, 101
  Federal Travel Regulation System                41, Subtitle F
  General                                         41, 300
  Payment From a Non-Federal Source for Travel    41, 304
       Expenses
  Payment of Expenses Connected With the Death    41, 303
       of Certain Employees
  Relocation Allowances                           41, 302
  Temporary Duty (TDY) Travel Allowances          41, 301
Geological Survey                                 30, IV
Government Accountability Office                  4, I
Government Ethics, Office of                      5, XVI
Government National Mortgage Association          24, III
Grain Inspection, Packers and Stockyards          7, VIII; 9, II
     Administration
Gulf Coast Ecosystem Restoration Council          40, VIII
Harry S. Truman Scholarship Foundation            45, XVIII
Health and Human Services, Department of          2, III; 5, XLV; 45, 
                                                  Subtitle A,
  Centers for Medicare & Medicaid Services        42, IV
  Child Support Enforcement, Office of            45, III
  Children and Families, Administration for       45, II, III, IV, X
  Community Services, Office of                   45, X
  Family Assistance, Office of                    45, II
  Federal Acquisition Regulation                  48, 3
  Food and Drug Administration                    21, I
  Human Development Services, Office of           45, XIII
  Indian Health Service                           25, V
  Inspector General (Health Care), Office of      42, V
  Public Health Service                           42, I
  Refugee Resettlement, Office of                 45, IV
Homeland Security, Department of                  2, XXX; 6, I; 8, I
  Coast Guard                                     33, I; 46, I; 49, IV
  Coast Guard (Great Lakes Pilotage)              46, III
  Customs and Border Protection                   19, I
  Federal Emergency Management Agency             44, I
  Human Resources Management and Labor Relations  5, XCVII
       Systems
  Immigration and Customs Enforcement Bureau      19, IV
  Transportation Security Administration          49, XII
HOPE for Homeowners Program, Board of Directors   24, XXIV
     of
Housing and Urban Development, Department of      2, XXIV; 5, LXV; 24, 
                                                  Subtitle B
  Community Planning and Development, Office of   24, V, VI
       Assistant Secretary for
  Equal Opportunity, Office of Assistant          24, I
       Secretary for
  Federal Acquisition Regulation                  48, 24
  Federal Housing Enterprise Oversight, Office    12, XVII
       of
  Government National Mortgage Association        24, III
  Housing--Federal Housing Commissioner, Office   24, II, VIII, X, XX
       of Assistant Secretary for
  Housing, Office of, and Multifamily Housing     24, IV
       Assistance Restructuring, Office of
  Inspector General, Office of                    24, XII
  Public and Indian Housing, Office of Assistant  24, IX
       Secretary for
  Secretary, Office of                            24, Subtitle A, VII
Housing--Federal Housing Commissioner, Office of  24, II, VIII, X, XX
     Assistant Secretary for
Housing, Office of, and Multifamily Housing       24, IV
     Assistance Restructuring, Office of
Human Development Services, Office of             45, XIII
Immigration and Customs Enforcement Bureau        19, IV
Immigration Review, Executive Office for          8, V
Independent Counsel, Office of                    28, VII

[[Page 1086]]

Indian Affairs, Bureau of                         25, I, V
Indian Affairs, Office of the Assistant           25, VI
     Secretary
Indian Arts and Crafts Board                      25, II
Indian Health Service                             25, V
Industry and Security, Bureau of                  15, VII
Information Resources Management, Office of       7, XXVII
Information Security Oversight Office, National   32, XX
     Archives and Records Administration
Inspector General
  Agriculture Department                          7, XXVI
  Health and Human Services Department            42, V
  Housing and Urban Development Department        24, XII, XV
Institute of Peace, United States                 22, XVII
Inter-American Foundation                         5, LXIII; 22, X
Interior Department                               2, XIV
  American Indians, Office of the Special         25, VII
       Trustee
  Bureau of Ocean Energy Management, Regulation,  30, II
       and Enforcement
  Endangered Species Committee                    50, IV
  Federal Acquisition Regulation                  48, 14
  Federal Property Management Regulations System  41, 114
  Fish and Wildlife Service, United States        50, I, IV
  Geological Survey                               30, IV
  Indian Affairs, Bureau of                       25, I, V
  Indian Affairs, Office of the Assistant         25, VI
       Secretary
  Indian Arts and Crafts Board                    25, II
  Land Management, Bureau of                      43, II
  National Indian Gaming Commission               25, III
  National Park Service                           36, I
  Natural Resource Revenue, Office of             30, XII
  Ocean Energy Management, Bureau of              30, V
  Reclamation, Bureau of                          43, I
  Secretary of the Interior, Office of            2, XIV; 43, Subtitle A
  Surface Mining Reclamation and Enforcement,     30, VII
       Office of
Internal Revenue Service                          26, I
International Boundary and Water Commission,      22, XI
     United States and Mexico, United States 
     Section
International Development, United States Agency   22, II
     for
  Federal Acquisition Regulation                  48, 7
International Development Cooperation Agency,     22, XII
     United States
International Joint Commission, United States     22, IV
     and Canada
International Organizations Employees Loyalty     5, V
     Board
International Trade Administration                15, III; 19, III
International Trade Commission, United States     19, II
Interstate Commerce Commission                    5, XL
Investment Security, Office of                    31, VIII
James Madison Memorial Fellowship Foundation      45, XXIV
Japan-United States Friendship Commission         22, XVI
Joint Board for the Enrollment of Actuaries       20, VIII
Justice Department                                2, XXVIII; 5, XXVIII; 28, 
                                                  I, XI; 40, IV
  Alcohol, Tobacco, Firearms, and Explosives,     27, II
       Bureau of
  Drug Enforcement Administration                 21, II
  Federal Acquisition Regulation                  48, 28
  Federal Claims Collection Standards             31, IX
  Federal Prison Industries, Inc.                 28, III
  Foreign Claims Settlement Commission of the     45, V
       United States
  Immigration Review, Executive Office for        8, V
  Offices of Independent Counsel                  28, VI
  Prisons, Bureau of                              28, V
  Property Management Regulations                 41, 128
Labor Department                                  5, XLII
  Employee Benefits Security Administration       29, XXV
  Employees' Compensation Appeals Board           20, IV
  Employment and Training Administration          20, V

[[Page 1087]]

  Employment Standards Administration             20, VI
  Federal Acquisition Regulation                  48, 29
  Federal Contract Compliance Programs, Office    41, 60
       of
  Federal Procurement Regulations System          41, 50
  Labor-Management Standards, Office of           29, II, IV
  Mine Safety and Health Administration           30, I
  Occupational Safety and Health Administration   29, XVII
  Office of Workers' Compensation Programs        20, VII
  Public Contracts                                41, 50
  Secretary of Labor, Office of                   29, Subtitle A
  Veterans' Employment and Training Service,      41, 61; 20, IX
       Office of the Assistant Secretary for
  Wage and Hour Division                          29, V
  Workers' Compensation Programs, Office of       20, I
Labor-Management Standards, Office of             29, II, IV
Land Management, Bureau of                        43, II
Legal Services Corporation                        45, XVI
Library of Congress                               36, VII
  Copyright Royalty Board                         37, III
  U.S. Copyright Office                           37, II
Local Television Loan Guarantee Board             7, XX
Management and Budget, Office of                  5, III, LXXVII; 14, VI; 
                                                  48, 99
Marine Mammal Commission                          50, V
Maritime Administration                           46, II
Merit Systems Protection Board                    5, II, LXIV
Micronesian Status Negotiations, Office for       32, XXVII
Military Compensation and Retirement              5, XCIV
     Modernization Commission
Millennium Challenge Corporation                  22, XIII
Mine Safety and Health Administration             30, I
Minority Business Development Agency              15, XIV
Miscellaneous Agencies                            1, IV
Monetary Offices                                  31, I
Morris K. Udall Scholarship and Excellence in     36, XVI
     National Environmental Policy Foundation
Museum and Library Services, Institute of         2, XXXI
National Aeronautics and Space Administration     2, XVIII; 5, LIX; 14, V
  Federal Acquisition Regulation                  48, 18
National Agricultural Library                     7, XLI
National Agricultural Statistics Service          7, XXXVI
National and Community Service, Corporation for   2, XXII; 45, XII, XXV
National Archives and Records Administration      2, XXVI; 5, LXVI; 36, XII
  Information Security Oversight Office           32, XX
National Capital Planning Commission              1, IV
National Commission for Employment Policy         1, IV
National Commission on Libraries and Information  45, XVII
     Science
National Council on Disability                    34, XII
National Counterintelligence Center               32, XVIII
National Credit Union Administration              5, LXXXVI; 12, VII
National Crime Prevention and Privacy Compact     28, IX
     Council
National Drug Control Policy, Office of           21, III
National Endowment for the Arts                   2, XXXII
National Endowment for the Humanities             2, XXXIII
National Foundation on the Arts and the           45, XI
     Humanities
National Highway Traffic Safety Administration    23, II, III; 47, VI; 49, V
National Imagery and Mapping Agency               32, I
National Indian Gaming Commission                 25, III
National Institute of Food and Agriculture        7, XXXIV
National Institute of Standards and Technology    15, II
National Intelligence, Office of Director of      32, XVII
National Labor Relations Board                    5, LXI; 29, I
National Marine Fisheries Service                 50, II, IV
National Mediation Board                          29, X
National Oceanic and Atmospheric Administration   15, IX; 50, II, III, IV, 
                                                  VI
National Park Service                             36, I

[[Page 1088]]

National Railroad Adjustment Board                29, III
National Railroad Passenger Corporation (AMTRAK)  49, VII
National Science Foundation                       2, XXV; 5, XLIII; 45, VI
  Federal Acquisition Regulation                  48, 25
National Security Council                         32, XXI
National Security Council and Office of Science   47, II
     and Technology Policy
National Telecommunications and Information       15, XXIII; 47, III, IV
     Administration
National Transportation Safety Board              49, VIII
Natural Resources Conservation Service            7, VI
Natural Resource Revenue, Office of               30, XII
Navajo and Hopi Indian Relocation, Office of      25, IV
Navy Department                                   32, VI
  Federal Acquisition Regulation                  48, 52
Neighborhood Reinvestment Corporation             24, XXV
Northeast Interstate Low-Level Radioactive Waste  10, XVIII
     Commission
Nuclear Regulatory Commission                     2, XX; 5, XLVIII; 10, I
  Federal Acquisition Regulation                  48, 20
Occupational Safety and Health Administration     29, XVII
Occupational Safety and Health Review Commission  29, XX
Ocean Energy Management, Bureau of                30, V
Offices of Independent Counsel                    28, VI
Office of Workers' Compensation Programs          20, VII
Oklahoma City National Memorial Trust             36, XV
Operations Office                                 7, XXVIII
Overseas Private Investment Corporation           5, XXXIII; 22, VII
Patent and Trademark Office, United States        37, I
Payment From a Non-Federal Source for Travel      41, 304
     Expenses
Payment of Expenses Connected With the Death of   41, 303
     Certain Employees
Peace Corps                                       2, XXXVII; 22, III
Pennsylvania Avenue Development Corporation       36, IX
Pension Benefit Guaranty Corporation              29, XL
Personnel Management, Office of                   5, I, XXXV; 45, VIII
  Human Resources Management and Labor Relations  5, XCVII
       Systems, Department of Homeland Security
  Federal Acquisition Regulation                  48, 17
  Federal Employees Group Life Insurance Federal  48, 21
       Acquisition Regulation
  Federal Employees Health Benefits Acquisition   48, 16
       Regulation
Pipeline and Hazardous Materials Safety           49, I
     Administration
Postal Regulatory Commission                      5, XLVI; 39, III
Postal Service, United States                     5, LX; 39, I
Postsecondary Education, Office of                34, VI
President's Commission on White House             1, IV
     Fellowships
Presidential Documents                            3
Presidio Trust                                    36, X
Prisons, Bureau of                                28, V
Privacy and Civil Liberties Oversight Board       6, X
Procurement and Property Management, Office of    7, XXXII
Productivity, Technology and Innovation,          37, IV
     Assistant Secretary
Public Contracts, Department of Labor             41, 50
Public and Indian Housing, Office of Assistant    24, IX
     Secretary for
Public Health Service                             42, I
Railroad Retirement Board                         20, II
Reclamation, Bureau of                            43, I
Recovery Accountability and Transparency Board    4, II
Refugee Resettlement, Office of                   45, IV
Relocation Allowances                             41, 302
Research and Innovative Technology                49, XI
     Administration
Rural Business-Cooperative Service                7, XVIII, XLII, L
Rural Development Administration                  7, XLII
Rural Housing Service                             7, XVIII, XXXV, L
Rural Telephone Bank                              7, XVI

[[Page 1089]]

Rural Utilities Service                           7, XVII, XVIII, XLII, L
Saint Lawrence Seaway Development Corporation     33, IV
Science and Technology Policy, Office of          32, XXIV
Science and Technology Policy, Office of, and     47, II
     National Security Council
Secret Service                                    31, IV
Securities and Exchange Commission                5, XXXIV; 17, II
Selective Service System                          32, XVI
Small Business Administration                     2, XXVII; 13, I
Smithsonian Institution                           36, V
Social Security Administration                    2, XXIII; 20, III; 48, 23
Soldiers' and Airmen's Home, United States        5, XI
Special Counsel, Office of                        5, VIII
Special Education and Rehabilitative Services,    34, III
     Office of
State Department                                  2, VI; 22, I; 28, XI
  Federal Acquisition Regulation                  48, 6
Surface Mining Reclamation and Enforcement,       30, VII
     Office of
Surface Transportation Board                      49, X
Susquehanna River Basin Commission                18, VIII
Technology Administration                         15, XI
Technology Policy, Assistant Secretary for        37, IV
Tennessee Valley Authority                        5, LXIX; 18, XIII
Thrift Supervision Office, Department of the      12, V
     Treasury
Trade Representative, United States, Office of    15, XX
Transportation, Department of                     2, XII; 5, L
  Commercial Space Transportation                 14, III
  Contract Appeals, Board of                      48, 63
  Emergency Management and Assistance             44, IV
  Federal Acquisition Regulation                  48, 12
  Federal Aviation Administration                 14, I
  Federal Highway Administration                  23, I, II
  Federal Motor Carrier Safety Administration     49, III
  Federal Railroad Administration                 49, II
  Federal Transit Administration                  49, VI
  Maritime Administration                         46, II
  National Highway Traffic Safety Administration  23, II, III; 47, IV; 49, V
  Pipeline and Hazardous Materials Safety         49, I
       Administration
  Saint Lawrence Seaway Development Corporation   33, IV
  Secretary of Transportation, Office of          14, II; 49, Subtitle A
  Surface Transportation Board                    49, X
  Transportation Statistics Bureau                49, XI
Transportation, Office of                         7, XXXIII
Transportation Security Administration            49, XII
Transportation Statistics Bureau                  49, XI
Travel Allowances, Temporary Duty (TDY)           41, 301
Treasury Department                               5, XXI; 12, XV; 17, IV; 
                                                  31, IX
  Alcohol and Tobacco Tax and Trade Bureau        27, I
  Community Development Financial Institutions    12, XVIII
       Fund
  Comptroller of the Currency                     12, I
  Customs and Border Protection                   19, I
  Engraving and Printing, Bureau of               31, VI
  Federal Acquisition Regulation                  48, 10
  Federal Claims Collection Standards             31, IX
  Federal Law Enforcement Training Center         31, VII
  Financial Crimes Enforcement Network            31, X
  Fiscal Service                                  31, II
  Foreign Assets Control, Office of               31, V
  Internal Revenue Service                        26, I
  Investment Security, Office of                  31, VIII
  Monetary Offices                                31, I
  Secret Service                                  31, IV
  Secretary of the Treasury, Office of            31, Subtitle A
  Thrift Supervision, Office of                   12, V
Truman, Harry S. Scholarship Foundation           45, XVIII
United States and Canada, International Joint     22, IV
     Commission
United States and Mexico, International Boundary  22, XI
   and Water Commission, United States Section
[[Page 1090]]

U.S. Copyright Office                             37, II
Utah Reclamation Mitigation and Conservation      43, III
     Commission
Veterans Affairs Department                       2, VIII; 38, I
  Federal Acquisition Regulation                  48, 8
Veterans' Employment and Training Service,        41, 61; 20, IX
     Office of the Assistant Secretary for
Vice President of the United States, Office of    32, XXVIII
Vocational and Adult Education, Office of         34, IV
Wage and Hour Division                            29, V
Water Resources Council                           18, VI
Workers' Compensation Programs, Office of         20, I
World Agricultural Outlook Board                  7, XXXVIII

[[Page 1091]]



List of CFR Sections Affected



All changes in this volume of the Code of Federal Regulations (CFR) that 
were made by documents published in the Federal Register since January 
1, 2010 are enumerated in the following list. Entries indicate the 
nature of the changes effected. Page numbers refer to Federal Register 
pages. The user should consult the entries for chapters, parts and 
subparts as well as sections for revisions.
For changes to this volume of the CFR prior to this listing, consult the 
annual edition of the monthly List of CFR Sections Affected (LSA). The 
LSA is available at www.fdsys.gov. For changes to this volume of the CFR 
prior to 2001, see the ``List of CFR Sections Affected, 1949-1963, 1964-
1972, 1973-1985, and 1986-2000'' published in 11 separate volumes. The 
``List of CFR Sections Affected 1986-2000'' is available at 
www.fdsys.gov.

                                  2010

14 CFR
                                                                   75 FR
                                                                    Page
Chapter I
13.19 (b) amended..................................................41979
    Regulation at 75 FR 41979 confirmed............................52859
13.27 (a) revised..................................................41979
    Regulation at 75 FR 41979 confirmed............................52859
13.305 (d) Tables 1, 2 and 3 removed; new table added..............72938
21 Special FAA conditions..........................................50688
    Airworthiness design standards.................................52860
    Regulation at 74 FR 53384 eff. date corrected to 4-16-11........9095
21.1 Regulation at 74 FR 53384 eff. date corrected to 4-16-11.......9095
21.2 Regulation at 74 FR 53384 eff. date corrected to 4-16-11.......9095
21.3 Regulation at 74 FR 53385 eff. date corrected to 4-16-11.......9095
21.5 Regulation at 74 FR 53385 eff. date corrected to 4-16-11.......9095
21.8 Regulation at 74 FR 53385 eff. date corrected to 4-16-11.......9095
21.9 Regulation at 74 FR 53385 eff. date corrected to 4-16-11.......9095
21.9 (b) and (c) introductory text corrected; eff. 4-16-11..........9095
21.15 Regulation at 74 FR 53385 eff. date corrected to 4-16-11......9095
21.20 Regulation at 74 FR 53385 eff. date corrected to 4-16-11......9095
21.21 Regulation at 74 FR 53385 eff. date corrected to 4-16-11......9095
21.27 Regulation at 74 FR 53386 eff. date corrected to 4-16-11......9095
21.29 Regulation at 74 FR 53386 eff. date corrected to 4-16-11......9095
21.33 Regulation at 74 FR 53386 eff. date corrected to 4-16-11......9095
21.45 Regulation at 74 FR 53386 eff. date corrected to 4-16-11......9095
21.47 Regulation at 74 FR 53386 eff. date corrected to 4-16-11......9095
21.50 Regulation at 74 FR 53386 eff. date corrected to 4-16-11......9095
21.53 Regulation at 74 FR 53386 eff. date corrected to 4-16-11......9095
21.73 Regulation at 74 FR 53387 eff. date corrected to 4-16-11......9095
21.75 Regulation at 74 FR 53387 eff. date corrected to 4-16-11......9095
21.97 Regulation at 74 FR 53387 eff. date corrected to 4-16-11......9095
21.113 Regulation at 74 FR 53387 eff. date corrected to 4-16-11.....9095
21.117 Regulation at 74 FR 53387 eff. date corrected to 4-16-11.....9095
21.119 Regulation at 74 FR 53387 eff. date corrected to 4-16-11.....9095
21.122 (a) corrected; eff. date corrected to 4-16-11................9095
21.122 Regulation at 74 FR 53387 eff. date corrected to 4-16-11.....9095

[[Page 1092]]

21.123 Regulation at 74 FR 53387 eff. date corrected to 4-16-11.....9095
21.125 Regulation at 74 FR 53387 eff. date corrected to 4-16-11.....9095
21.130 Regulation at 74 FR 53387 eff. date corrected to 4-16-11.....9095
21.131--21.150 (Subpart G) Regulation at 74 FR 53387 eff. date 
        corrected to 4-16-11........................................9095
21.231 Regulation at 74 FR 53390 eff. date corrected to 4-16-11.....9095
21.251 Regulation at 74 FR 53390 eff. date corrected to 4-16-11.....9095
21.253 Regulation at 74 FR 53394 eff. date corrected to 4-16-11.....9095
21.267 Regulation at 74 FR 53394 eff. date corrected to 4-16-11.....9095
21.271 Regulation at 74 FR 53394 eff. date corrected to 4-16-11.....9095
21.293 Regulation at 74 FR 53394 eff. date corrected to 4-16-11.....9095
21.301--21.320 (Subpart K) Regulation at 74 FR 53394 eff. date 
        corrected to 4-16-11........................................9095
21.601--21.621 (Subpart O) Regulation at 74 FR 53392 eff. date 
        corrected to 4-16-11........................................9095
21.621 Heading correctly revised; eff. date corrected to 4-16-11 
                                                                    9095
23 Special FAA conditions..............20516, 20518, 44881, 50850, 50853
25 Special FAA conditions.......32, 35, 37, 39, 2433, 2434, 6092, 12965, 
         18399, 26643, 27926, 38391, 46838, 46840, 47176, 52614, 70090, 
                                                            70092, 71346
    Regulation at 74 FR 26946 corrected.............................1527
25.571 (a)(3) introductory text and (b) introductory text revised; 
        eff. 1-14-11...............................................69781
25.1322 Revised; eff. 1-3-11.......................................67209
25.1583 (a)(3) revised.............................................49818
25 Appendix H amended; eff. 1-14-11................................69782
26.5 Revised; eff. 1-14-11.........................................69782
26.21--26.23 (Subpart C) added; eff. 1-14-11.......................69782
26.43 (e) introductory text revised................................11734
26.45 (b)(1) and (e)(1) revised....................................11734
27.1457 (d)(1)(ii) revised.........................................17045
27.1459 (a)(3)(ii) revised.........................................17045
29 Special FAA conditions...................................50688, 77524
29.1457 (d)(1)(ii) revised.........................................17045
29.1459 (a)(3)(ii) revised.........................................17045
39.13........222, 225, 902, 905, 908, 913, 1018, 1528, 1531, 1534, 1537, 
                                     1539, 1699, 2056, 2059, 2061, 2064,
2066, 2068, 2788, 3126, 3135, 3142, 3146, 3148, 3151, 3616, 3984, 4267, 
4476, 4479, 4481, 4484, 4486, 5678, 5683, 5685, 5687, 5690, 5692, 5694, 
5696, 6850, 6853, 6856, 7028, 7344, 7930, 7932, 7935, 7937, 7939, 7941, 
7944, 7946, 7948, 7950, 8462, 8466, 8469, 8472, 8475, 8478, 8480, 9516, 
9754, 9757, 9762, 10661, 10665, 10668, 10670, 11424, 11430, 11434, 
11437, 11441, 12439, 12440, 12442, 12658, 12660, 12663, 12664, 12667, 
12669, 12673, 12969, 12971, 14335, 15322, 15324, 15327, 15330, 15334, 
16647, 16650, 16653, 16656, 16659, 16661, 16663, 16665, 17296, 19195, 
19198, 19202, 19205, 19200, 19208, 19211, 20266, 21162, 21501, 22504, 
22507, 22509, 22512, 22513, 22516, 22518, 22520, 22523, 22694, 23569, 
23572, 23576, 23578, 24391, 26882, 26884, 26887, 27403, 27405, 27407, 
27410, 27412, 27415, 27418, 27420, 27423, 27425, 28464, 28467, 28471, 
28473, 28477, 28479, 28481, 28484, 28486, 28752, 30269, 30271, 30273, 
30276, 30278, 30281, 30283, 30285, 30288, 30292, 30293, 30688, 32091, 
32252, 32254, 32259, 32261, 32262, 32265, 32267, 32650, 33160, 33163, 
34348, 34351, 34356, 34358, 34925, 35606, 35610, 35612, 35614, 35617, 
35621, 35623, 35626, 37710, 37991, 37992, 37996, 38003, 38000, 38008, 
38010, 38013, 38016, 38019, 38395, 38400, 38405, 39144, 39788, 39792, 
39797, 39799, 39802, 39804, 39810, 39813, 39816, 39819, 42586, 42590, 
42595, 43051, 43396, 43802, 43805, 43808, 43811, 47181, 47184, 47186, 
47193, 47196, 47198, 47200, 47203, 47207, 47209, 49367, 49369, 49372, 
49376, 49378, 50856, 50858, 50862, 50864, 50866, 50868, 50870, 50873, 
50876, 50878, 50879, 51653, 51655, 51657, 51658, 51660, 51933, 52234, 
52236, 52239, 52241, 52244, 52248, 52252, 52254, 52257, 52265, 52436, 
53844, 53847, 53850, 53853, 53856, 53859, 53861, 53863, 54463, 55454, 
55457, 55460, 55462, 57372, 57660, 57662, 57665, 57668, 57845, 58291, 
59062, 59064, 59066, 59070, 59072, 59075,

[[Page 1093]]

59078, 59607, 60303, 60603, 60606, 60609, 60612, 60616, 61047, 61339, 
61342, 61344, 61346, 61350, 61353, 61976, 61978, 61981, 61983, 61986, 
61988, 61991, 62320, 63040, 63042, 63044, 63047, 63049, 63052, 63054, 
63057, 63059, 63061, 63063, 63065, 64113, 64634, 64638, 65223, 66650, 
66652, 66654, 66657, 66658, 66660, 67612, 67614, 68171, 68173, 68175, 
68178, 68181, 68183, 68187, 68683, 68685, 68687, 68689, 68691, 68694, 
68696, 68699, 69859, 69861, 69864, 70097, 70099, 70102, 70103, 70105, 
70107, 70110, 70814, 71352, 71529, 71531, 71533, 71535, 71537, 71539, 
71542, 78598
    Corrected.....9515, 13225, 22693, 23573, 23580, 28188, 31282, 47208, 
                         53846, 57846, 65224, 69862, 71353, 78883, 80293
    Eff. date corrected............................................38019
    Technical correction...........................................43397
    Eff. 1-3-11....................................................72654
    Eff. 1-5-11...............................74609, 74612, 74619, 79951
    Eff. 1-11-11.......................75869, 75872, 75877, 75880, 75883
    Eff. 1-12-11............................................81426, 81435
    Eff. 1-13-11...................................................76625
    Eff. 1-14-11...................................................82222
    Eff. 1-20-11..............................78590, 78592, 78595, 78600
    Eff. 1-21-11...................................................78882
    Eff. 1-25-11...................................................79953
    Eff. 2-1-11..........81411, 81416, 81418, 81421, 81424, 81429, 81432
43.1 (b) revised....................................................5219
43.2 Regulation at 74 FR 53394 eff. date corrected to 4-16-11.......9095
43.3 Regulation at 74 FR 53394 eff. date corrected to 4-16-11.......9095
45 Regulation at 74 FR 53394 eff. date corrected to 4-16-11.........9095
45.1 Regulation at 74 FR 53394 eff. date corrected to 4-16-11.......9095
45.10--45.16 (Subpart B) Regulation at 74 FR 53394 eff. date 
        corrected to 4-16-11........................................9095
45.10 Regulation at 74 FR 53394 eff. date corrected to 4-16-11......9095
45.14 Regulation at 74 FR 53395 eff. date corrected to 4-16-11......9095
45.15 Regulation at 74 FR 53395 eff. date corrected to 4-16-11......9095
45.16 Regulation at 74 FR 53395 eff. date corrected to 4-16-11......9095
47 Nomenclature change.............................................41979
    Regulation at 75 FR 41979 confirmed............................52859
    Technical correction...........................................58292
47.1 Revised.......................................................41979
    Regulation at 75 FR 41979 confirmed............................52859
47.2 Amended.......................................................41979
    Regulation at 75 FR 41979 confirmed............................52859
47.3 (a) revised; (b)(2) amended...................................41979
    Regulation at 75 FR 41979 confirmed............................52859
47.5 (a) and (c) amended...........................................41979
    Regulation at 75 FR 41979 confirmed............................52859
47.7 (a), (c)(2) introductory text and (d) introductory text 
        amended....................................................41979
    (d) introductory text amended..................................41980
    Regulation at 75 FR 41979 and 41980 confirmed..................52859
47.8 (c) amended............................................41979, 41980
    Regulation at 75 FR 41979 and 41980 confirmed..................52859
47.9 (a) introductory text and (e) amended.........................41979
    Regulation at 75 FR 41979 confirmed............................52859
47.11 Introductory text, (e) and (h) amended.......................41979
    (b)(1) amended.................................................41980
    Regulation at 75 FR 41979 and 41980 confirmed..................52859
47.13 (a) through (f) revised......................................41980
    Regulation at 75 FR 41980 confirmed............................52859
47.15 Amended; (d) amended; undesignated text following (a)(3) 
        designated as (a)(4); (a) introductory text, (2), new (4), 
        (c) and (f) revised; (i) and (j) added.....................41980
    Regulation at 75 FR 41980 confirmed............................52859
47.16 (a) amended..................................................41981
    Regulation at 75 FR 41981 confirmed............................52859
47.17 (a)(4), (5) and (6) revised; (a)(7) added....................41981
    Regulation at 75 FR 41981 confirmed............................52859
47.31 (a) introductory text amended................................41979
    (a)(1) and (c) revised; (a)(2) amended; (d) removed............41981

[[Page 1094]]

    Regulation at 75 FR 41979 and 41981 confirmed..................52859
47.33 (b) and (d) amended..........................................41979
    (a)(2) revised; (c) amended....................................41981
    Regulation at 75 FR 41979 and 41981 confirmed..................52859
47.35 (a) introductory text and (b) amended........................41979
    Regulation at 75 FR 41979 confirmed............................52859
47.37 (a)(2) amended...............................................41979
    Regulation at 75 FR 41979 confirmed............................52859
47.39 Revised......................................................41981
    Regulation at 75 FR 41981 confirmed............................52859
47.40 Added........................................................41981
    Regulation at 75 FR 41981 confirmed............................52859
47.41 (a) introductory text and (b)(3) revised; (a)(2) and (4) 
        removed; (a)(3) and (5) through (9) redesignated as new 
        (a)(2) through (7); (a)(1) through (5) amended; (b)(4) 
        added......................................................41982
    Regulation at 75 FR 41982 confirmed............................52859
47.43 (b) amended..................................................41979
    (b) revised....................................................41982
    Regulation at 75 FR 41979 and 41982 confirmed..................52859
47.45 Revised......................................................41982
    Regulation at 75 FR 41982 confirmed............................52859
47.47 (a) introductory text and (1) revised........................41982
    Regulation at 75 FR 41982 confirmed............................52859
47.49 Revised......................................................41982
    Regulation at 75 FR 41982 confirmed............................52859
47.51 Removed......................................................41983
    Regulation at 75 FR 41983 confirmed............................52859
47.61 Heading, (a) introductory text and (2) revised; (b) amended; 
        (c) added..................................................41983
    Regulation at 75 FR 41983 confirmed............................52859
47.63 (a) amended..................................................41983
    Regulation at 75 FR 41983 confirmed............................52859
47.65 Revised......................................................41983
    Regulation at 75 FR 41983 confirmed............................52859
47.67 Revised......................................................41983
    Regulation at 75 FR 41983 confirmed............................52859
47.69 Introductory text amended....................................41983
    Regulation at 75 FR 41983 confirmed............................52859
47.71 (a) amended; (b) revised.....................................41983
    Regulation at 75 FR 41983 confirmed............................52859

                                  2011

14 CFR
                                                                   76 FR
                                                                    Page
Chapter I
1.1 Amended............................................................9
    Regulation at 76 FR 9 eff. date confirmed.......................9495
    Regulation at 73 FR 76213 OMB number changed...................34576
17 Revised.........................................................55221
21.35 (b)(2) revised...............................................64233
21.231--21.293 (Subpart J) Removed..................................8893
21.431--21.493 (Subpart M) Removed..................................8893
21.700 (Subpart P) Added; interim..................................12555
23 Special FAA conditions.......17332, 19903, 45399, 52230, 65101, 72087
23.3 (d) amended; eff. 1-31-12.....................................75753
23.45 (h) introductory text revised; eff. 1-31-12..................75753
23.49 Heading, (a) introductory text and (c) introductory text 
        revised; eff. 1-31-12......................................75753
23.51 (b)(1) introductory text and (c) introductory text revised; 
        eff. 1-31-12...............................................75753
23.53 (c) revised; eff. 1-31-12....................................75753
23.55 Introductory text revised; eff. 1-31-12......................75753
23.57 Introductory text revised; eff. 1-31-12......................75753
23.59 Introductory text revised; eff. 1-31-12......................75753
23.61 Introductory text revised; eff. 1-31-12......................75753
23.63 (c) introductory text and (d) introductory text revised; 
        eff. 1-31-12...............................................75753
23.65 (b) revised; eff. 1-31-12....................................75753
23.67 (c) redesignated as (d); (b) introductory text, (1) 
        introductory text and new (d) introductory text revised; 
        new (c) added; eff. 1-31-12................................75754

[[Page 1095]]

23.73 Revised; eff. 1-31-12........................................75754
23.77 (b) introductory text and (c) introductory text revised; 
        eff. 1-31-12...............................................75754
23.177 (a), (b) and (d) revised; eff. 1-31-12......................75754
23.181 (b) revised; eff. 1-31-12...................................75755
23.201 (d) and (e) revised; (e) redesignated as (f); new (e) 
        added; eff. 1-31-12........................................75755
23.203 (c) revised; eff. 1-31-12...................................75755
23.251 Revised; eff. 1-31-12.......................................75755
23.253 (b)(1) and (2) revised; (b)(3) and (d) added; eff. 1-31-12 
                                                                   75755
23.255 Added; eff. 1-31-12.........................................75755
23.561 (e)(1) added; eff. 1-31-12..................................75756
23.562 (a) introductory text, (b) introductory text and (c)(5)(ii) 
        revised; eff. 1-31-12......................................75756
23.571 (d) added; eff. 1-31-12.....................................75756
23.629 (b)(1), (3), (4) and (c) revised; eff. 1-31-12..............75756
23.703 Introductory text revised; (c) added; eff. 1-31-12..........75757
23.735 (e) revised; eff. 1-31-12...................................75757
23.777 (d) revised; eff. 1-31-12...................................75757
23.807 (e)(3) added; eff. 1-31-12..................................75757
23.831 (c) and (d) added; eff. 1-31-12.............................75757
23.841 (a) and (b)(6) revised; (c) and (d) added; eff. 1-31-12.....75757
23.853 (d)(2) revised; eff. 1-31-12................................75759
23.856 Added; eff. 1-31-12.........................................75759
23.903 (b)(3) added; eff. 1-31-12..................................75759
23.1165 (f) revised; eff. 1-31-12..................................75759
23.1193 (g) revised; eff. 1-31-12..................................75759
23.1195 (a) introductory text and (2) revised; eff. 1-31-12........75759
23.1197 Introductory text revised; eff. 1-31-12....................75760
23.1199 Introductory text revised; eff. 1-31-12....................75760
23.1201 Introductory text revised; eff. 1-31-12....................75760
23.1301 (b) and (c) revised; (d) removed; eff. 1-31-12.............75760
23.1303 (c) revised; eff. 1-31-12..................................75760
23.1306 Added......................................................33135
23.1309 Revised; eff. 1-31-12......................................75760
23.1310 Added; eff. 1-31-12........................................75760
23.1311 (a)5, (6), (7) and (b) revised; eff. 1-31-12...............75760
23.1323 (e) revised; eff. 1-31-12..................................75761
23.1331 (c) revised; eff, 1-31-12..................................75761
23.1353 (h) revised; eff. 1-31-12..................................75761
23.1431 (a) revised; eff. 1-31-12..................................75761
23.1443 Revised; eff. 1-31-12......................................75761
23.1445 (c) added; eff. 1-31-12....................................75762
23.1447 (g) added; eff. 1-31-12....................................75762
23.1505 (c) revised; eff. 1-31-12..................................75762
23.1545 (d) revised; eff. 1-31-12..................................75762
23.1555 (d)(3) added; eff. 1-31-12.................................75763
23.1559 (d) added; eff. 1-31-12....................................75763
23.1563 (d) added; eff. 1-31-12....................................75763
23.1567 (e) added; eff. 1-31-12....................................75763
23.1583 (c)(3) introductory text, (4) introductory text and (5) 
        introductory text revised; (c)(4)(iii) and (iv) 
        redesignated as (c)(4)(ii)(A) and (B); eff. 1-31-12........75763
23.1585 (f) introductory text revised; eff. 1-31-12................75763
23.1587 (d) introductory text revised; eff. 1-31-12................75763
23 Appendix F amended; eff. 1-31-12................................75763
    Appendix C corrected; CFR correction...........................81790
25 Special FAA conditions........5061, 8278, 10213, 10476, 10482, 14794, 
         14795, 15798, 17022, 17334, 25229, 30523, 31223, 31451, 31453, 
         31454, 31456, 35324, 35736, 36861, 36863, 36864, 36865, 36870, 
         38550, 39763, 41041, 41045, 44245, 44246, 54923, 57625, 57627, 
                  62603, 63818, 63822, 63823, 65103, 65105, 71865, 81360
25.21 (g)(1) revised; eff. 1-30-12.................................74654
25.107 (e)(1)(iv) revised; eff. 1-30-12............................74654
25.177 Revised; eff. 1-30-12.......................................74654
25.253 (a)(4) and (5) added; (b) introductory text and (c) 
        introductory text revised; eff. 1-30-12....................74654
25.509 (a)(3)(ii) corrected; CFR correction........................81790
25.1316 Revised....................................................33135
25.1801 (Subpart I) Added; interim.................................12555
27 Special FAA conditions............................10489, 12274, 20490
27.573 Added; eff. 1-30-12.........................................74663
27.610 (d)(4) revised..............................................33135
27.1309 (d) removed................................................33135
27.1316 Added......................................................33135
27 Appendix B amended..............................................33135
    Appendix A amended; eff. 1-30-12...............................74663
29.571 Revised; eff. 1-31-12.......................................75442
29.573 Added; eff. 1-30-12.........................................74664
29.610 (d)(4) revised..............................................33135

[[Page 1096]]

29.1309 (h) removed................................................33136
29.1316 Added......................................................33136
29 Appendix A amended; eff. 1-30-12................................74663
33 Special FAA conditions.....................30819, 33981, 40594, 56097
    Policy statement...............................................47423
    Technical correction...........................................55553
33.27 Revised......................................................42023
34 Policy statement................................................44457
39 Policy statement.................................................9495
39.13........254, 420, 422, 425, 427, 429, 431, 433, 436, 438, 443, 445, 
            1341, 1345, 1348, 1350, 1353, 1982, 1984, 1987, 1992, 1995, 
            1997, 4057, 4218, 4220, 4223, 4225, 4227, 5067, 5468, 6324, 
            6524, 6527, 6532, 6534, 6536, 6537, 6540, 6542, 6546, 6550, 
            7102, 8606, 8608, 8611, 8614, 8617, 8619, 8621, 8623, 9500, 
          9964, 10216, 10218, 10221, 10225, 10229, 10232, 11326, 11942, 
         12278, 12557, 12846, 13060, 13062, 13064, 13066, 13068, 13071, 
         13073, 13075, 13077, 13079, 13081, 14797, 14799, 15802, 15804, 
         15806, 15812, 15815, 15819, 15821, 15824, 16527, 17758, 17760, 
         18021, 18024, 18027, 18030, 18032, 18035, 18039, 18377, 18868, 
         20230, 20232, 20495, 20497, 20499, 20502, 20505, 22006, 22008, 
         22297, 22300, 22304, 22307, 22310, 22315, 22318, 22320, 23170, 
         24795, 24797, 24801, 25536, 27221, 27228, 27234, 27238, 27241, 
         27243, 27245, 27247, 27862, 27865, 27868, 27874, 27877, 27881, 
         28170, 28631, 28632, 28635, 28637, 28639, 30531, 24345, 24347, 
         24350, 24352, 24355, 24357, 24359, 24361, 31458, 31461, 31463, 
         31466, 31798, 31800, 31802, 31818, 33983, 33986, 33987, 33989, 
         33992, 35329, 35333, 35335, 35338, 35341, 35343, 35345, 35348, 
         36284, 36982, 37242, 37246, 37249, 37253, 37254, 37256, 37259, 
         37631, 39250, 39253, 39256, 39258, 40218, 40220, 40223, 41396, 
         41649, 41652, 41655, 41658, 41661, 41664, 41666, 41668, 41672, 
         41675, 42026, 42030, 42032, 42035, 44254--44259, 44461, 45656, 
         45659, 47057, 47426, 47429, 47432, 50112, 50114, 50116, 50404, 
         50406, 50883, 52215, 52216, 52218, 52221, 52223, 52228, 53047, 
         53302, 53304, 53306, 53310, 53314, 53316, 53320, 53325, 53327, 
         53634, 54094, 54374, 54927, 55782, 55784, 55787, 56278, 56281, 
         56285, 56288, 56291, 56638, 57632, 57901, 58096, 58099, 59009, 
         59012, 59241, 60368, 60712, 61034, 61037, 61257, 61556, 61559, 
         61560, 61562, 63157, 63160, 63162, 63166, 63168, 63170, 63174, 
         63178, 64003, 64783, 64786, 64789, 64792, 64794, 64796, 64799, 
         64802, 65937, 65940, 65943, 66607, 66611, 66614, 66616, 66618, 
         66620, 66622, 66625, 67053, 67342, 67345, 67348, 67592, 67596, 
         68298, 68300, 68303, 68305, 68309, 68636, 68638, 69124, 70041, 
         70043, 70045, 70048, 70335, 70337, 71243, 71247, 72093, 73488, 
           73490, 74666, 75444, 75773, 76294, 77381, 77383, 79495, 81792
    Corrected.......255, 7695, 25535, 27239, 29998, 30254, 36981, 40596, 
                                       46598, 59013, 62605, 64003, 69123

[[Page 1097]]

    Eff. 1-3-12.................73479, 73482, 73485, 73492, 73497, 78140
    Eff. 1-5-12....................................................74668
    Eff. 1-10-12...................................................76028
    Eff. 1-17-12........77368, 77370, 77372, 77376, 77377, 77379, 77108, 
                                                                   77110
    Eff. 1-23-12..............................78521, 78523, 78525, 78528
    Eff. 1-24-12............................................78807, 78809
    Eff. 1-25-12...................................................79053
    Eff. 1-27-12...................................................80229
    Eff. 2-3-12...............................82105, 82107, 82111, 82112
45.11 (g)(1)(ii) and (3) amended....................................7486

                                  2012

14 CFR
                                                                   77 FR
                                                                    Page
Chapter I
1.1 Amended.........................................................9166
    Regulation at 77 FR 9166 eff. date delayed.....................22186
    Regulation at 77 FR 9166 confirmed......................39388, 40478
    Correctly amended..............................................62147
21 Policy statement.........................................11385, 38463
21.97 (a)(2) revised; eff. 2-4-13..................................71695
21.101 (a), (b) introductory text, (3) and (c) revised; eff. 2-4-
        13.........................................................71695
23 Special FAA conditions............................16907, 17323, 42949
25 Special FAA conditions...1618, 2437, 5990, 6945, 16910, 17325, 18099, 
         21861, 32006, 36123, 38467, 40255, 56525, 57481, 64023, 64025, 
                                                     64029, 67251, 67557
    Policy statement...............................................11385
25.729 (a)(1)(ii), (iii), (b), (e) introductory text, (5), (f) 
        introductory text and (1) revised; (e)(7) and (f)(3) added
                                                                    1617
25.773 (b)(2) revised; (b)(3) and (4) added.........................1618
26.21 Table 1 correctly revised....................................30878
27 Technical correction.............................................4890
    Special FAA conditions.........................................48058
27.573 Regulation at 76 FR 74663 confirmed.........................50576
27 Regulation at 76 FR 74663 confirmed.............................50576
29 Technical correction.............................................4890
    Special FAA conditions..................................44110, 60883
29.571 Regulation at 76 FR 75442 confirmed.........................50576
29.573 Regulation at 76 FR 74664 confirmed.........................50576
29 Regulation at 76 FR 74664 confirmed.............................50576
33 Technical correction.............................................2015
33.83 (a) revised..................................................39624
    (a) correctly revised..........................................58301
33.87 (a)(1), (b)(1), (2) heading, (i), (ii), (3), (4), (5), 
        (c)(1), (d)(1), (e)(1) and (2) amended.....................22187
34 Authority citation revised......................................76849
34.1 Amended.......................................................76849
34.2 Amended.......................................................76850
34.3 (c) and (d) revised...........................................76850
34.7 (b) amended; (d) revised......................................76850
34.9 Added.........................................................76850
34.10 Revised......................................................76850
34.21 (b), (d), (e) and (f) revised; (g) added.....................76851
34.23 Added........................................................76851
34.31 (b) and (c) revised..........................................76852
34.48 Added........................................................76852
34.60 Revised......................................................76853
34.61 Removed......................................................76854
34.62 Removed......................................................76854
34.63 Removed......................................................76854
34.64 Removed......................................................76854
34.71 Removed......................................................76854
34.80--34.89 (Subpart H) Removed...................................76854
39 Technical correction............................................12991
39.13....2, 4, 730, 734, 1010, 1623, 1625, 1627, 2441, 2445, 3089, 3091, 
            3381, 3384, 3582, 3584, 3586, 3588, 4218, 4648, 4649, 4651, 
                                                       5168, 5993, 5995,
 5997, 5999, 6001, 6004, 6664, 6667, 6669, 6670, 6673, 7520, 7522, 7524, 
8093, 8723, 9168, 9519, 9521, 9838, 10354, 10356, 10951, 12159, 12164, 
12168, 12172, 12174, 12176, 12178, 12181, 12447, 12450, 12452, 12990, 
13190, 13192, 13194, 13484, 13487, 13489, 14680, 14683, 15940, 16137, 
16139, 16141, 16145, 16147, 16150, 16154, 16156, 16429, 16431, 16433, 
16663, 16915, 16917, 16919, 16920, 16922, 19060, 19062, 19064, 19066, 
19068, 19070, 19073, 19075, 20507, 20510, 20514, 20517, 20519, 20521, 
20524, 20527, 21396, 21399, 21401, 21403, 21418, 21421, 21425, 21428, 
21432, 22189, 23111, 23381, 23383, 23386, 23389, 24343, 24346, 24348, 
24351, 24352, 24354, 24357, 24359, 24361, 24364, 24366, 24368, 24586, 
24830, 24832, 24834, 24837, 26155, 26157, 26160, 26665, 26938, 26944, 
26947, 26949, 28240, 28242, 29209, 29211, 29213, 29215, 29856, 29859, 
29862, 29864, 30186, 30374, 30882, 31168, 31170, 31173, 31175, 31177,

[[Page 1098]]

31485, 32008, 32010, 32886, 32888, 32891, 32893, 33084, 33621, 33623, 
34207, 36126, 36128, 36130, 36133, 36136, 36138, 36141, 36145, 36148, 
36390, 37284, 38003, 37767, 37769, 37772, 37774, 37776, 37779, 37781, 
37783, 37785, 37787, 37789, 37792, 37794, 37796, 37798, 38470, 38471, 
39155, 39156, 39158, 39160, 39626, 40480, 40483, 40487, 41044, 41047, 
41888, 41891, 41893, 41896, 41899, 42420, 42422, 42425, 42953, 42956, 
42957, 42960, 42964, 42968, 42972, 44115, 44117, 44119, 44431, 44434, 
44436, 44438, 46930, 46934, 46936, 46939, 46942, 46945, 46947, 47271, 
47275, 47276, 47278, 48420, 48422, 48424, 48426, 48428, 49704, 49707, 
49709, 49711, 50372, 50580, 50583, 51461, 51463, 51696, 51893, 52202, 
52204, 52208, 52215, 54354, 54788, 54792, 54795, 54797, 54800, 54802, 
54804, 55412, 55682, 55686, 55688, 56530, 56756, 56757, 56761, 56988, 
56990, 56992, 57000, 57003, 57004, 57007, 57009, 57485, 57991, 57995, 
57997, 57999, 58003, 58005, 58472, 58762, 58764, 58766, 58926, 59727, 
59730, 59734, 60287, 60291, 60300, 60888, 60890, 60893, 61512, 63216, 
63712, 63713, 63715, 63717, 64694, 64696, 64699, 64701, 64703, 64705, 
64708, 64710, 64713, 65612, 65614, 65616, 65618, 65800, 65802, 65804, 
65806, 65809, 65811, 65814, 67255, 67260, 67262, 67266, 67269, 67560, 
67764, 67766, 68051, 68054, 68056, 68058, 68060, 68062, 68064, 69557, 
69560, 69741, 69743, 69746, 69751, 70115, 70356, 70359, 70361, 70364, 
70367, 70370, 71086, 71088, 72202, 72204, 73269, 73275, 74581
    Corrected......1009, 5387, 10952, 12992, 17327, 20700, 20987, 24137, 
                                              42625, 57995, 66535, 67562
    Eff. 1-7-13........................71484, 71486, 71488, 71490, 71492
    Eff. 1-10-13............................................75829, 75832
    Eff. 1-11-13...................................................72914
    Eff. 1-14-13..............................73267, 73271, 73281, 73286
    Eff. 1-16-13..............................73901, 73905, 73908, 73910
    Eff. 1-30-13............................................75826, 75834
    Eff. 1-31-13.....................................76229, 76231, 76233
43.3 (k) added; eff. 1-28-13.......................................71096
43 Appendix A amended; eff. 1-28-13................................71096
45.13 (a)(7) introductory text revised; (a)(7)(iv) and (v) added 
                                                                   76854

                                  2013

14 CFR
                                                                   78 FR
                                                                    Page
Chapter I
Chapter I Policy statement..................................36412, 42419
16 Authority citation revised......................................56141
16.1 (a) introductory text and (3) through (6) revised.............56141
16.3 Amended.......................................................56141
16.11 Heading, (a) and (b) introductory text revised; (c) and (d) 
        added......................................................56142
16.13 (a) through (d) and (f) revised; (h) and (i) added...........56142
16.15 (a), (b), (d)(1) and (2) revised; (d)(3) added...............56143
16.17 (c) revised..................................................56143
16.19 (d) and (e) added............................................56143
16.21 Revised......................................................56143
16.23 Heading, (a), (b)(2), (4), (c), (d) and (j) revised; (k) and 
        (l) added..................................................56144
16.25 Revised......................................................56144
16.26 Added........................................................56144
16.27 Revised......................................................56145
16.29 (b)(2) amended...............................................56145
16.31 Revised......................................................56145
16.33 Revised......................................................56145
16.34 Added........................................................56145
16.105 Amended.....................................................56146
16.109 Revised.....................................................56146
16.201 (b) revised.................................................56146
16.203 (a)(1), (b)(1) and (2) revised..............................56146
16.207 Revised.....................................................56146
16.211 (c) amended.................................................56147
16.215 (e) added...................................................56147
16.227 Revised.....................................................56147
16.229 Introductory text added.....................................56147
16.233 Revised.....................................................56147
16.235 (b) revised.................................................56147
16.241 Transferred to Subpart F; (a) and (c) revised; (f) removed 
                                                                   56147
16.243 Transferred to Subpart F....................................56147
16.241--16.243 (Subpart G) Removed.................................56147
16.245 Added.......................................................56147
16.247 (Subpart G) Redesignated from Subpart H.....................56148

[[Page 1099]]

16.247 (Subpart H) Redesignated as Subpart G.......................56148
16.247 (a), (b)(2) and (4) revised.................................56148
16.301--16.307 (Subpart H) Redesignated from Subpart I.............56148
16.301--16.307 (Subpart I) Redesignated as new Subpart H...........56148
16.301 Removed; new section redesignated from 16.303...............56148
16.303 Redesignated as 16.301; redesignated from 16.305............56148
16.305 Redesignated as 16.303; redesignated from 16.307............56148
16.307 Redesignated as 16.305......................................56148
21 Policy statement..................................16779, 50313, 68687
21.93 (b)(5) added..................................................1138
23 Special FAA conditions......10055, 28719, 35747, 36084, 37958, 50317, 
                                                            55629, 57470
25 Special FAA conditions.........6195, 6198, 8961, 11089, 11553, 11554, 
         11555, 11556, 11560, 11562, 14005, 14007, 14155, 19981, 21037, 
         25846, 31835, 31836, 31838, 32078, 36084, 41684, 49655, 63845, 
         63847, 63848, 65153, 65155, 65156, 67291, 68985, 68986, 70848, 
                  70849, 73993, 73995, 75451, 75453, 76734, 76736, 76980
    Special FAA conditions; eff. 1-21-14...........................76731
25.1302 Added......................................................25846
29 Special FAA conditions..........................................35108
33 Special FAA conditions...................................15597, 19982
34 Authority citation revised......................................63016
    Comment disposition............................................63017
34.1 Amended.......................................................63016
34.23 (a)(2) and (b)(1) revised....................................63017
    Table corrected................................................65554
35.15 (c) and (d) revised...........................................4041
    (c) and (d) correctly revised..................................45052
35.16 Added.........................................................4042
36.1 (a)(5) and (i) added; (c) amended..............................1138
36.13 Added.........................................................1139
36.1101--36.1103 (Subpart K) Added..................................1139
36 Appendix K added.................................................1139
39.13..............................................................6, 8,
12, 16, 858, 1725, 1728, 1729, 1731, 1732, 1734, 1737, 1741, 2196, 2198, 
2199, 2332, 4044, 4049, 4052, 4055, 4056, 4760, 4763, 5127, 5711, 5713, 
6201, 6204, 6207, 6726, 7258, 7260, 7262, 7263, 7642, 7644, 7646, 7649, 
9310, 9583, 9778, 9780, 9782, 9784, 9787, 9789, 9791, 9793, 9795, 9797, 
9800, 9801, 10500, 10502, 11564, 11566, 11568, 11571, 11573, 11974, 
11977, 11979, 12232, 14159, 14162, 14163, 14165, 14443, 14641, 14643, 
14647, 14649, 15278, 15280, 15282, 15601, 15872, 15875, 16606, 17072, 
17074, 17076, 17078, 17079, 17081, 17083, 17288, 17293, 17296, 17298, 
17592, 17594, 17866, 18225, 18227, 18229, 18231, 19086, 19089, 19091, 
19095, 19984, 20228, 20232, 20236, 21228, 21232, 21234, 21238, 22167, 
22169, 22171, 22174, 22177, 22179, 22182, 22184, 22187, 23106, 23109, 
23111, 23113, 23115, 23460, 23464, 23467, 23470, 23832, 24040, 24043, 
24339, 24340, 24343, 24673, 24676, 25362, 25364, 25366, 25369, 25371, 
25373, 25376, 25379, 25381, 26235, 26242, 27003, 27006, 27012, 27017, 
27022, 28127, 28129, 28131, 28724, 28727, 28729, 28731, 31387, 31392, 
31395, 32082, 32346, 32348, 32351, 32354, 32552, 33196, 33198, 33200, 
33202, 33205, 33207, 34551, 35113, 35749, 35751, 35754, 36090, 36409, 
37439, 37447, 37450, 37702, 37705, 38196, 38545, 38547, 38551, 38553, 
38823, 38825, 38827, 39568, 39573, 39575, 40955, 40958, 41275, 41279, 
41282, 41284, 41286, 41288, 42409, 42411, 42414, 42416, 42418, 42678, 
43762, 43765, 43768, 43769, 43771, 44424, 44872, 45053, 45055, 47528, 
47530, 47533, 47536, 47539, 47545, 47547, 47552, 48288, 48600, 48796, 
49110, 49112, 49114, 49661, 49663, 49905, 49907, 49909, 49912, 49914, 
49916, 50321, 51049, 51052, 51054, 51056, 51060, 52406, 52409, 52411, 
52413, 52415, 52418, 52420, 52837, 52840, 52842, 53237, 53632, 53634, 
53637, 53639, 53641, 54151, 54155, 54379, 54382, 54384, 54386, 54388, 
54752, 56149, 56590, 56593, 56596, 56598, 56600, 56603, 57048, 57051, 
57054, 57255, 57785, 57787, 58870, 58873, 58875, 59225, 59801, 59804, 
60184, 60186, 60188, 60190, 60657, 60659, 60663, 60669, 60672, 60675, 
60678, 60680, 60683, 61163, 61167, 61170, 61173, 61176, 61179, 63852, 
63855, 63857, 63859, 64158, 64160, 64163, 64166, 64396, 65162, 65164, 
65167, 65170, 65172, 65174, 65177, 65180, 65182, 65184, 65186, 65188, 
65191, 65194, 65196, 65199, 65202,

[[Page 1100]]

65204, 65205, 65208, 65870, 65872, 66253, 66256, 66260, 67011, 67013, 
67015, 67017, 67019, 67022, 67023, 68347, 68350, 68353, 68356, 68359, 
68690, 68692, 68696, 68698, 69989, 69991, 70197, 70199, 70201, 70204, 
70206, 70208, 70210, 70213, 70215, 70217, 70488, 70491, 70852, 71991, 
73688, 76047
    Corrected.....2615, 13463, 24985, 29613, 41837, 49115, 49116, 54561, 
                                                     56151, 63853, 68360
    Corrected; eff. 1-14-14........................................79599
    Eff. 1-6-14......................................71994, 71997, 72000
    Eff. 1-7-14...72551, 72553, 72557, 72560, 72563, 72566, 72568, 72569
    Eff. 1-8-14....................................................72792
    Eff. 1-13-14...................................................73690
    Eff. 1-14-14............................................73999, 79294
    Eff. 1-21-14..............................76037, 76041, 76049, 76051
    Eff. 1-24-14...................................................76985
    Eff. 1-28-14.....................................77566, 77568, 77570
    Eff. 1-31-14................78696, 78700, 78702, 78704, 78707, 78711
    Eff. 2-3-14......................................79289, 79291, 79297
45 Comment disposition.............................................63017

                                  2014

14 CFR
                                                                   79 FR
                                                                    Page
Chapter I
Policy statement...................................................66282
11 Authority citation revised......................................12937
11.201 (b) table amended...........................................12937
13 Authority citation revised......................................46967
    Regulation at 79 FR 46967 confirmed............................61761
13.20 (c) and (d) revised..........................................46967
    Regulation at 79 FR 46967 confirmed............................61761
23 Special FAA conditions..........................................26111
25 Special FAA conditions......1591, 2359, 2365, 7054, 7370, 7372, 9379, 
          9380, 11679, 20768, 32633, 32635, 32636, 32637, 32639, 32640, 
         32642, 33043, 33669, 33673, 33674, 33675, 33677, 34215, 41418, 
         41419, 41633, 43232, 43233, 43236, 43237, 43239, 44657, 44658, 
         46167, 46169, 46170, 46171, 46173, 49423, 49426, 49427, 49429, 
         52165, 52169, 53128, 53129, 54571, 54572, 54574, 54576, 56485, 
           57429, 58243, 60319, 63299, 63300, 63302, 65562, 68107, 73469
    Technical correction....................................21845, 59431
25.21 (g)(1) and (2) revised; (g)(3) and (4) added; eff. 1-5-15....65524
25.105 (a)(2) introductory text revised; eff. 1-5-15...............65525
25.111 (c)(5)(i) and (ii) revised; eff. 1-5-15.....................65525
25.119 (b) revised; eff. 1-5-15....................................65525
25.121 (b)(2)(ii) introductory text, (c)(2)(ii) introductory text 
        and (d)(2)(ii) revised; eff. 1-5-15........................65525
25.123 (b)(2) introductory text revised; eff. 1-5-15...............65525
25.125 (a)(2), (b)(2)(ii)(B) and (C) revised; eff. 1-5-15..........65525
25.143 (c) introductory text, (i)(1) and (j) introductory text 
        revised; eff. 1-5-15.......................................65525
25.207 (b), (e)(1) through (5) and (h) introductory text revised; 
        eff. 1-5-15................................................65526
25.237 (a)(3)(ii) revised; eff. 1-5-15.............................65526
25.253 (c) introductory text revised; eff. 1-5-15..................65526
25.307 (a) revised.................................................59429
25.331 (c) introductory text and (2) revised; eff. 2-9-15..........73466
25.341 (a)(5)(i), (6) and (b) revised; (c) added; eff. 2-9-15......73467
25.343 (b)(1)(ii) revised; eff. 2-9-15.............................73468
25.345 (c)(2) revised; eff. 2-9-15.................................73468
25.361 Revised; eff. 2-9-15........................................73468
25.362 Added; eff. 2-9-15..........................................73468
25.371 Revised; eff. 2-9-15........................................73468
25.373 (a) revised; eff. 2-9-15....................................73468
25.391 Introductory text revised; eff. 2-9-15......................73468
25.395 (b) revised; eff. 2-9-15....................................73468
25.415 Revised; eff. 2-9-15........................................73468
25.621 (a), (c) and (d) revised....................................59429
25.683 Revised.....................................................59430
25.721 Revised.....................................................59430
25.773 (b)(1)(ii) revised; eff. 1-5-15.............................65526
25.787 (a) revised.................................................59430
25.795 (d) and (e) redesignated as (e) and (f); new (d) added......13519
25.903 (a)(3) added; eff. 1-5-15...................................65526
25.929 (a) revised; eff. 1-5-15....................................65526
25.963 (d) revised.................................................59430
25.994 Revised.....................................................59431
25.1093 (b) revised; eff. 1-5-15...................................65526
25.1323 (i) revised; eff. 1-5-15...................................65527
25.1324 Added; eff. 1-5-15.........................................65527
25.1325 (b) revised; eff. 1-5-15...................................65527
25.1420 Added; eff. 1-5-15.........................................65528
25.1450 (b)(3) added...............................................13519

[[Page 1101]]

25.1517 Revised; eff. 2-9-15.......................................73469
25.1521 (c)(3) redesignated as (c)(4) and revised; new (c)(3) 
        added; eff. 1-5-15.........................................65528
25.1533 (c) added; eff. 1-5-15.....................................65528
25 Appendix C amended; Appendix O added; eff. 1-5-15...............65528
    Appendix G removed; eff. 2-9-15................................73469
27 Special FAA conditions............................47553, 48946, 48949
29 Special FAA conditions............................54889, 75423, 78694
33 Special FAA conditions..........................................64666
    Authority citation revised.....................................65536
33.68 Revised; eff. 1-5-15.........................................65536
33.77 (a) added; (c) introductory text, (1), (d) and (e) revised; 
        eff. 1-5-15................................................65537
33 Appendix D added; eff. 1-5-15...................................65538
36 Technical correction............................................20769
36.1 (h)(5) redesignated as (h)(7); new (h)(5) and (6) added; new 
        (h)(7) revised.............................................12044
36.11 (c) revised; (d) added.......................................12044
36.805 (b)(1) and (2) revised......................................12045
36 Appendices H and J amended......................................12045
39.13...............................................................345,
534, ;538, 541, 544, 548, 550, 1316, 1734, 2367, 3304, 3482, 4268, 4272, 
5249, 5251, 5253, 5256, 5258, 7375, 7379, 7382, 7384, 7388, 7389, 8082, 
9384, 9386, 9388, 9391, 9394, 9397, 9399, 9401, 9991, 9993, 10332, 
10334, 10336, 10339, 10341, 10343, 10346, 10348, 10349, 10352, 10960, 
11688, 11692, 11694, 11696, 11699, 11700, 11702, 12047, 12365, 12367, 
12370, 12372, 12375, 12377, 13198, 13200, 13203, 13205, 13207, 13520, 
13523, 13525, 13528, 13530, 13532, 14170, 14979, 14985, 14991, 14995, 
14998, 15653, 15657, 15659, 15664, 15666, 15669, 15672, 15675, 16174, 
17010, 17392, 17394, 17396, 17398, 17400, 17405, 17407, 17410, 17419, 
17858, 18613, 18616, 18618, 18621, 18624, 18627, 18630, 18989, 19813, 
21387, 21388, 21390, 21393, 21846, 22363, 22365, 22368, 22370, 23261, 
23895, 23899, 23901, 23905, 23908, 23910, 23912, 24543, 24548, 24550, 
24552, 24555, 24557, 26604, 26607, 26610, 26611, 27482, 27486, 30006, 
30013, 30017, 31850, 31853, 31856, 31858, 32861, 33047, 33049, 33051, 
33053, 33056, 33680, 34404, 34407, 35034, 35036, 35039, 35480, 35483, 
35487, 36389, 36631, 36633, 36635, 36637, 36639, 37168, 37171, 37173, 
39301, 39957, 39959, 39961, 39962, 41086, 41089, 41092, 41094, 41097, 
41100, 41103, 41107, 41110, 41113, 41116, 41119, 41122, 42648, 42651, 
42654, 42657, 42659, 42662, 42664, 43606, 43610, 43612, 43615, 43618, 
44269, 44662, 44665, 44668, 44671, 44675, 44678, 45086, 45319, 45324, 
45326, 45328, 45331, 45334, 45336, 45339, 45342, 46970, 48020, 48023, 
48027, 48029, 48032, 48953, 48956, 48959, 48961, 48964, 48967, 48971, 
48974, 49433, 49436, 49441, 49444, 49447, 49451, 50540, 50544, 51073, 
51076, 51079, 51082, 51085, 51089, 51230, 51233, 51236, 51239, 51241, 
52174, 52176, 52179, 52182, 52186, 52189, 52192, 52547, 52955, 53287, 
53289, 54579, 54582, 54892, 54894, 54896, 54899, 55605, 56267, 57752, 
57755, 59092, 59095, 59101, 59103, 59632, 59634, 59638, 59642, 59644, 
60323, 60326, 60328, 60330, 60333, 60336, 60338, 60342, 63307, 63310, 
63313, 63504, 63811, 64083, 64085, 64088, 64091, 64093, 64096, 64306, 
64308, 64506, 66281, 67043, 67046, 67050, 67054, 67341, 67345, 68110, 
69034, 69036, 69038, 69040, 69371, 70441, 70443, 70446, 70450, 70456, 
70460, 70787, 71306
    Eff. 1-2-15....................................................70789
    Eff. 1-6-15...............................71298, 71301, 71303, 71309
    Eff. 1-9-15...............................72123, 72126, 72129, 72134
    Eff. 1-13-15...................................................72969
    Eff. 1-16-15................73802, 73804, 73807, 73811, 73813, 73816
    Eff. 1-20-15..............................74598, 74601, 74604, 74606
    Eff. 1-28-15............................................77375, 77382
    Eff. 1-27-15...................................................77377
    Corrected......................................................77384
43.10 (c)(6) correctly revised.....................................67055


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