[Federal Register Volume 70, Number 69 (Tuesday, April 12, 2005)]
[Proposed Rules]
[Pages 19015-19027]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 05-7320]


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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 25

[Docket No. NM305; Notice No. 25-05-04-SC]


Special Conditions: Airbus Model A380-800 Airplane; Dynamic 
Braking, Interaction of Systems and Structures, Limit Pilot Forces, 
Side Stick Controllers, Dive Speed Definition, Electronic Flight 
Control System-Lateral-Directional Stability, Longitudinal Stability, 
and Low Energy Awareness, Electronic Flight Control System-Control 
Surface Awareness, Electronic Flight Control System-Flight 
Characteristics Compliance Via the Handling Qualities Rating Method, 
Flight Envelope Protection-General Limiting Requirements, Flight 
Envelope Protection-Normal Load Factor (G) Limiting, Flight Envelope 
Protection-High Speed Limiting, Flight Envelope Protection-Pitch and 
Roll Limiting, Flight Envelope Protection-High Incidence Protection and 
Alpha-Floor Systems, High Intensity Radiated Fields (HIRF) Protection, 
and Operation Without Normal Electrical Power

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Notice of proposed special conditions.

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SUMMARY: This notice proposes special conditions for the Airbus A380-
800 airplane. This airplane will have novel or unusual design features 
when compared to the state of technology envisioned in the 
airworthiness standards for transport category airplanes. These design 
features include side stick controllers, a body landing gear in 
addition to conventional wing and nose landing gears, electronic flight 
control systems, and flight envelope protection. These proposed special 
conditions also pertain to the effects of such novel or unusual design 
features, such as their effects on the structural performance of the 
airplane. Finally, the proposed special conditions pertain to the 
effects of certain conditions on these novel or unusual design 
features, such as the effects of high intensity radiated fields (HIRF) 
or of operation without normal electrical power. Additional special 
conditions will be issued for other novel or unusual design features of 
the Airbus A380-800 airplanes. A list is provided in the section of 
this document entitled ``Discussion of Novel or Unusual Design 
Features.''

DATES: Comments must be received on or before May 27, 2005.

ADDRESSES: Comments on this proposal may be mailed in duplicate to: 
Federal Aviation Administration, Transport Airplane Directorate, 
Attention: Rules Docket (ANM-113), Docket No. NM305, 1601 Lind Avenue 
SW., Renton, Washington 98055-4056; or delivered in duplicate to the 
Transport Airplane Directorate at the above address. All comments must 
be marked: Docket No. NM305. Comments may be inspected in the Rules 
Docket weekdays, except Federal holidays, between 7:30 a.m. and 4 p.m.

FOR FURTHER INFORMATION CONTACT: Holly Thorson, FAA, International 
Branch, ANM-116, Transport Airplane Directorate, Aircraft Certification 
Service, 1601 Lind Avenue SW., Renton, Washington 98055-4056; telephone 
(425) 227-1357; facsimile (425) 227-1149.

SUPPLEMENTARY INFORMATION:

Comments Invited

    The FAA invites interested persons to participate in this 
rulemaking by submitting written comments, data, or views. The most 
helpful comments reference a specific portion of the special 
conditions, explain the reason for any recommended change, and include 
supporting data. We ask that you send us two copies of written 
comments.
    We will file in the docket all comments we receive as well as a 
report summarizing each substantive public contact with FAA personnel 
concerning these proposed special conditions. The docket is available 
for public inspection before and after the comment closing date. If you 
wish to review the docket in person, go to the address in the ADDRESSES 
section of this notice between 7:30 a.m. and 4 p.m., Monday through 
Friday, except Federal holidays.
    We will consider all comments we receive on or before the closing 
date for comments. We will consider comments filed late, if it is 
possible to do so without incurring expense or delay. We may change the 
proposed special conditions in light of the comments we receive.
    If you want the FAA to acknowledge receipt of your comments on this 
proposal, include with your comments a pre-addressed, stamped postcard 
on which the docket number appears. We will stamp the date on the 
postcard and mail it back to you.

Background

    Airbus applied for FAA certification/validation of the 
provisionally-designated Model A3XX-100 in its letter AI/L 810.0223/98, 
dated August 12, 1998, to the FAA. Application for certification by the 
Joint Aviation Authorities (JAA) of Europe had been made on January 16, 
1998, reference AI/L 810.0019/98. In its letter to the FAA, Airbus 
requested an extension to the 5-year period for type certification in 
accordance with 14 CFR 21.17(c). The request was for an extension to a 
7-year period, using the date of the initial application letter to the 
JAA as the reference date. The reason given by Airbus for the request 
for extension is related to the technical challenges, complexity, and 
the number of new and novel features on the airplane. On November 12, 
1998, the Manager, Aircraft Engineering Division, AIR-100, granted 
Airbus' request for the 7-year period based on the date of application 
to the JAA.
    In its letter AI/LE-A 828.0040/99 Issue 3, dated July 20, 2001, 
Airbus stated that its target date for type certification of the Model 
A380-800 has been moved from May 2005, to January 2006, in order to 
match the delivery date of the first production airplane. In accordance 
with 14 CFR 21.17(d)(2), Airbus chose a new application date of April 
20, 1999, and requested that the 7-year certification period which had 
already been approved be continued. The part 25 certification basis for 
the Model A380-800 airplane was adjusted to reflect the new application 
date.
    The Model A380-800 airplane will be an all-new, four-engine jet 
transport airplane with a full double-deck, two-aisle cabin. The 
maximum takeoff weight will be 1.235 million pounds with a typical 
three-class layout of 555 passengers.

Type Certification Basis

    Under the provisions of 14 CFR 21.17, Airbus must show that the 
Model A380-800 airplane meets the applicable provisions of 14 CFR part 
25, as amended by Amendments 25-1 through 25-98. If the Administrator 
finds that the applicable airworthiness regulations do not contain 
adequate or appropriate safety standards for the Airbus A380-800 
airplane because of novel or unusual design features, special 
conditions are prescribed under the provisions of 14 CFR 21.16.
    In addition to the applicable airworthiness regulations and special

[[Page 19016]]

conditions, the Airbus Model A380-800 airplane must comply with the 
fuel vent and exhaust emission requirements of 14 CFR part 34 and the 
noise certification requirements of 14 CFR part 36. In addition, the 
FAA must issue a finding of regulatory adequacy pursuant to section 611 
of Public Law 93-574, the ``Noise Control Act of 1972.''
    Special conditions, as defined in 14 CFR 11.19, are issued in 
accordance with 14 CFR 11.38 and become part of the type certification 
basis in accordance with 14 CFR 21.17(a)(2), Amendment 21-69, effective 
September 16, 1991.
    Special conditions are initially applicable to the model for which 
they are issued. Should the type certificate for that model be amended 
later to include any other model that incorporates the same novel or 
unusual design feature, or should any other model already included on 
the same type certificate be modified to incorporate the same novel or 
unusual design features, the special conditions would also apply to the 
other model under the provisions of 14 CFR 21.101(a)(1), Amendment 21-
69, effective September 16, 1991.

Discussion of Novel or Unusual Design Features

    The Airbus A380-800 airplane will incorporate a number of novel or 
unusual design features. Because of rapid improvements in airplane 
technology, the applicable airworthiness regulations do not contain 
adequate or appropriate safety standards for these design features. The 
special conditions proposed for Airbus Model A380 contain the 
additional safety standards that the Administrator considers necessary 
to establish a level of safety equivalent to that established by the 
existing airworthiness standards.
    These proposed special conditions are identical or nearly identical 
to those previously required for type certification of the basic Model 
A340 airplane or earlier models. One exception is the special condition 
pertaining to Interaction of Systems and Structures. It was not 
required for the basic Model A340 but was required for type 
certification of the larger, heavier Model A340-500 and -600 airplanes.
    In general, the proposed special conditions were derived initially 
from standardized requirements developed by the Aviation Rulemaking 
Advisory Committee (ARAC), comprised of representatives of the FAA, 
Europe's Joint Aviation Authorities (now replaced by the European 
Aviation Safety Agency), and industry. In some cases, a draft Notice of 
Proposed Rulemaking has been prepared but no final rule has yet been 
promulgated.
    Additional special conditions will be issued for other novel or 
unusual design features of the Airbus Model A380-800 airplane. Those 
proposed special conditions pertain to the following topics:
     Fire protection,
     Evacuation, including availability of stairs in an 
emergency,
     Emergency exit arrangement--outside viewing,
     Escape system inflation systems,
     Escape systems installed in non-pressurized compartments,
     Ground turning loads,
     Crashworthiness,
     Flotation and ditching,
     Discrete gust requirements,
     Transient engine failure loads,
     Airplane jacking loads,
     Landing gear pivoting loads,
     Design roll maneuvers, and
     Extendable length escape systems.

1. Dynamic Braking

    The A380 landing gear system will include body gear in addition to 
the conventional wing and nose gear. This landing gear configuration 
may result in more complex dynamic characteristics than those found in 
conventional landing gear configurations. Section 25.493(d) by itself 
does not contain an adequate standard for assessing the braking loads 
for the A380 landing gear configuration.
    Due to the potential complexities of the A380 landing gear system, 
in addition to meeting the requirements of Sec.  25.493(d), a rational 
analysis of the braked roll conditions is necessary. Airbus Model A340-
500 and -600 also have a body-mounted main landing gear in addition to 
the wing and nose gears. Therefore, a special condition similar to that 
required for that model is appropriate for the model A380-800.

2. Interaction of Systems and Structures

    The A380 is equipped with systems which affect the airplane's 
structural performance either directly or as a result of failure or 
malfunction. The effects of these systems on structural performance 
must be considered in the certification analysis. This analysis must 
include consideration of normal operation and of failure conditions 
with required structural strength levels related to the probability of 
occurrence.
    Previously, special conditions have been specified to require 
consideration of the effects of systems on structures. The special 
condition proposed for the Model A380 is nearly identical to that 
issued for the Model A340-500 and -600 series airplanes.

3. Limit Pilot Forces

    Like some other Airbus models, the Model A380 airplane is equipped 
with a side stick controller instead of a conventional control stick. 
This kind of controller is designed to be operated using only one hand. 
The requirement of Sec.  25.397(c), which defines limit pilot forces 
and torques for conventional wheel or stick controls, is not 
appropriate for a side stick controller. Therefore, a special condition 
is necessary to specify the appropriate loading conditions for this 
kind of controller.
    A special condition for side stick controllers has already been 
developed for the Airbus model A320 and A340 airplanes, both of which 
also have a side stick controller instead of a conventional control 
stick. The same special condition would be appropriate for the model 
A380 airplane.

4. Side Stick Controllers

    The A380--like its predecessors, the A320, A330, and A340--will use 
side stick controllers for pitch and roll control. Regulatory 
requirements for conventional wheel and column controllers, such as 
requirements pertaining to pilot strength and controllability, are not 
directly applicable to side stick controllers. In addition, pilot 
control authority may be uncertain, because the side sticks are not 
mechanically interconnected as with conventional wheel and column 
controls.
    In previous Airbus airplane certification programs, special 
conditions pertaining to side stick controllers were addressed in three 
separate issue papers, entitled ``Pilot Strength,'' ``Pilot Coupling,'' 
and ``Pilot Control.'' The resulting separate special conditions are 
combined in this special condition under the title of ``Side Stick 
Controllers.'' In order to harmonize with the JAA, the following has 
been added to Special Condition 4.c. Side Stick Controllers:
    Pitch and roll control force and displacement sensitivity must be 
compatible, so that normal inputs on one control axis will not cause 
significant unintentional inputs on the other.

5. Dive Speed Definition

    Airbus proposes to reduce the speed spread between VC 
and VD required by Sec.  25.335(b), based on the 
incorporation of a high speed protection system in the A380 flight 
control laws. The A380--like the A320, A330, and A340--is equipped with 
a high speed protection system which limits nose down pilot

[[Page 19017]]

authority at speeds above VC/MC and prevents the 
airplane from actually performing the maneuver required under Sec.  
25.335(b)(1).
    Section 25.335(b)(1) is an analytical envelope condition which was 
originally adopted in Part 4b of the Civil Air Regulations to provide 
an acceptable speed margin between design cruise speed and design dive 
speed. Freedom from flutter and airframe design loads is affected by 
the design dive speed. While the initial condition for the upset 
specified in the rule is 1g level flight, protection is afforded for 
other inadvertent overspeed conditions as well. Section 25.335(b)(1) is 
intended as a conservative enveloping condition for all potential 
overspeed conditions, including non-symmetric ones. To establish that 
all potential overspeed conditions are enveloped, the applicant should 
demonstrate either of the following:
     Any reduced speed margin--based on the high speed 
protection system in the A380--will not be exceeded in inadvertent or 
gust induced upsets, resulting in initiation of the dive from non-
symmetric attitudes; or
     The airplane is protected by the flight control laws from 
getting into non-symmetric upset conditions.
    In addition, the high speed protection system in the A380 must have 
a high level of reliability.

6. Electronic Flight Control System: Lateral-Directional Stability, 
Longitudinal Stability, and Low Energy Awareness

    In lieu of compliance with the regulations pertaining to lateral-
directional and longitudinal stability, this special condition ensures 
that the model A380 will have suitable airplane handling qualities 
throughout the normal flight envelope (reference paragraphs 6.a. and 
6.b.).
    The unique features of the A380 flight control system and side-
stick controllers, when compared with conventional airplanes with wheel 
and column controllers, do not provide conventional awareness to the 
flight crew of a change in speed or a change in the direction of flight 
(reference paragraph 6.c.). This special condition requires that 
adequate awareness be provided to the pilot of a low energy state (low 
speed, low thrust, and low altitude) below normal operating speeds.
    a. Lateral-Directional Static Stability: The model A380 airplane 
has a flight control design feature within the normal operational 
envelope in which side stick deflection in the roll axis commands roll 
rate. As a result, the stick force in the roll axis will be zero 
(neutral stability) during the straight, steady sideslip flight 
maneuver of Sec.  25.177(c) and will not be ``substantially 
proportional to the angle of sideslip,'' as required by the regulation.
    The electronic flight control system (EFCS) on the A380 as on its 
predecessors--the A320, A330 and A340--contains fly-by-wire control 
laws that result in neutral lateral-directional static stability. 
Therefore, the conventional requirements of the regulations are not 
met.
    With conventional control system requirements, positive static 
directional stability is defined as the tendency to recover from a skid 
with the rudder free. Positive static lateral stability is defined as 
the tendency to raise the low wing in a sideslip with the aileron 
controls free. The regulations are intended to accomplish the 
following:
     Provide additional cues of inadvertent sideslips and skids 
through control force changes.
     Ensure that short periods of unattended operation do not 
result in any significant changes in yaw or bank angle.
     Provide predictable roll and yaw response.
     Provide acceptable level of pilot attention (i.e., 
workload) to attain and maintain a coordinated turn.
    b. Longitudinal Static and Dynamic Stability: The longitudinal 
flight control laws for the A380 provide neutral static stability 
within the normal operational envelope. Therefore, the airplane design 
does not comply with the static longitudinal stability requirements of 
Sec. Sec.  25.171, 25.173, and 25.175.
    Static longitudinal stability on conventional airplanes with 
mechanical links to the pitch control surface means that a pull force 
on the controller will result in a reduction in speed relative to the 
trim speed, and a push force will result in higher than trim speed. 
Longitudinal stability is required by the regulations for the following 
reasons:
     Speed change cues are provided to the pilot through 
increased and decreased forces on the controller.
     Short periods of unattended control of the airplane do not 
result in significant changes in attitude, airspeed or load factor.
     A predictable pitch response is provided to the pilot.
     An acceptable level of pilot attention (i.e., workload) to 
attain and maintain trim speed and altitude is provided to the pilot.
     Longitudinal stability provides gust stability.
    The pitch control movement of the side stick is a normal load 
factor or ``g'' command which results in an initial movement of the 
elevator surface to attain the commanded load factor. That movement is 
followed by integrated movement of the stabilizer and elevator to 
automatically trim the airplane to a neutral (1g) stick-free stability. 
The flight path commanded by the initial side stick input will remain 
stick-free until the pilot gives another command. This control function 
is applied during ``normal'' control law within the speed range from 
Vaprot (the speed at the angle of attack protection limit) 
to VMO to MMO. Once outside this speed range, the 
control laws introduce the conventional longitudinal static stability 
as described above.
    As a result of neutral static stability, the A380 does not meet the 
requirements of part 25 for static longitudinal stability.
    c. Low Energy Awareness: Static longitudinal stability provides an 
awareness to the flight crew of a low energy state (low speed and 
thrust at low altitude). Past experience on airplanes fitted with a 
flight control system which provides neutral longitudinal stability 
shows there are insufficient feedback cues to the pilot of excursion 
below normal operational speeds. The maximum angle of attack protection 
system limits the airplane angle of attack and prevents stall during 
normal operating speeds, but this system is not sufficient to prevent 
stall at low speed excursions below normal operational speeds. Until 
intervention, there are no stability cues, because the airplane remains 
trimmed. Additionally, feedback from the pitching moment due to thrust 
variation is reduced by the flight control laws. Recovery from a low 
speed excursion may become hazardous when the low speed is associated 
with low altitude and the engines are operating at low thrust or with 
other performance limiting conditions.

7. Electronic Flight Control System: Control Surface Awareness

    With a response-command type of flight control system and no direct 
coupling from cockpit controller to control surface, such as on the 
A380, the pilot is not aware of the actual surface deflection position 
during flight maneuvers. Some unusual flight conditions, arising from 
atmospheric conditions or airplane or engine failures or both, may 
result in full or nearly full surface deflection. Unless the flight 
crew is made aware of excessive deflection or impending control surface 
deflection limiting, piloted or auto-flight system control of the 
airplane might be inadvertently continued in a way which would cause 
loss of control or other

[[Page 19018]]

unsafe handling or performance characteristics.
    This special condition requires that suitable annunciation be 
provided to the flight crew when a flight condition exists in which 
nearly full control surface deflection occurs. Suitability of such a 
display must take into account that some pilot-demanded maneuvers 
(e.g., rapid roll) are necessarily associated with intended full or 
nearly full control surface deflection. Therefore, simple alerting 
systems which would function in both intended or unexpected control-
limiting situations must be properly balanced between needed crew 
awareness and not getting nuisance warnings.

8. Electronic Flight Control System: Flight Characteristics Compliance 
Via the Handling Qualities Rating Method (HQRM)

    The Model A380 airplane will have an Electronic Flight Control 
System (EFCS). This system provides an electronic interface between the 
pilot's flight controls and the flight control surfaces (for both 
normal and failure states). The system also generates the actual 
surface commands that provide for stability augmentation and control 
about all three airplane axes. Because EFCS technology has outpaced 
existing regulations--written essentially for unaugmented airplanes 
with provision for limited ON/OFF augmentation--suitable special 
conditions and a method of compliance are required to aid in the 
certification of flight characteristics.
    This special condition and the method of compliance presented in 
Appendix 7 of the Flight Test Guide, AC 25-7A, provide a means by which 
one may evaluate flight characteristics--as, for example, 
``satisfactory,'' ``adequate,'' or ``controllable''--to determine 
compliance with the regulations. The HQRM in Appendix 7 was developed 
for airplanes with control systems having similar functions and is 
employed to aid in the evaluation of the following:
     All EFCS/airplane failure states not shown to be extremely 
improbable and where the envelope (task) and atmospheric disturbance 
probabilities are each 1.
     All combinations of failures, atmospheric disturbance 
level, and flight envelope not shown to be extremely improbable.
    The HQRM provides a systematic approach to the assessment of 
handling qualities. It is not intended to dictate program size or need 
for a fixed number of pilots to achieve multiple opinions. The airplane 
design itself and success in defining critical failure combinations 
from the many reviewed in Systems Safety Assessments would dictate the 
scope of any HQRM application.
    Handling qualities terms, principles, and relationships familiar to 
the aviation community have been used to formulate the HQRM. For 
example, we have established that the well-known COOPER-HARPER rating 
scale and the proposed FAA three-part rating system are similar. This 
approach is derived in part from the contract work on the flying 
qualities of highly augmented/ relaxed static stability airplanes, in 
relation to regulatory and flight test guide requirements. The work is 
reported in DOT/FAA/CT-82/130, Flying Qualities of Relaxed Static 
Stability Aircraft, Volumes I and II.

9. Flight Envelope Protection: General Limiting Requirements

    This special condition and the following ones--pertaining to flight 
envelope protection--present general limiting requirements for all the 
unique flight envelope protection features of the basic A380 Electronic 
Flight Control System (EFCS) design. Current regulations do not address 
these types of protection features. The general limiting requirements 
are necessary to ensure a smooth transition from normal flight to the 
protection mode and adequate maneuver capability. The general limiting 
requirements also ensure that the structural limits of the airplane are 
not exceeded. Furthermore, failure of the protection feature must not 
create hazardous flight conditions. Envelope protection parameters 
include angle of attack, normal load factor, bank angle, pitch angle, 
and speed. To accomplish these envelope protections, one or more 
significant changes occur in the EFCS control laws as the normal flight 
envelope limit is approached or exceeded.
    Each specific type of envelope protection is addressed individually 
in the special conditions which follow.

10. Flight Envelope Protection: Normal Load Factor (G) Limiting

    The A380 flight control system design incorporates normal load 
factor limiting on a full time basis that will prevent the pilot from 
inadvertently or intentionally exceeding the positive or negative 
airplane limit load factor. This limiting feature is active in all 
normal and alternate flight control modes and cannot be overridden by 
the pilot. There is no requirement in the regulations for this limiting 
feature.
    Except for the Airbus airplanes with fly-by-wire flight controls, 
the normal load factor limit is unique in that traditional airplanes 
with conventional flight control systems (mechanical linkages) are 
limited in the pitch axis only by the elevator surface area and 
deflection limit. The elevator control power is normally derived for 
adequate controllability and maneuverability at the most critical 
longitudinal pitching moment. The result is that traditional airplanes 
have a significant portion of the flight envelope in which 
maneuverability in excess of limit structural design values is 
possible.
    Part 25 does not require a demonstration of maneuver control or 
handling qualities beyond the design limit structural loads. 
Nevertheless, some pilots have become accustomed to the availability of 
this excess maneuver capacity in case of extreme emergency, such as 
upset recoveries or collision avoidance. Airbus is aware of the concern 
and has published the results of its research which indicate the 
following:
     Pilots rarely, if ever, use the excess maneuvering 
capacity in collision avoidance maneuvers, and
     Other features of its flight control system would have 
prevented most, if not all, of the upset cases on record where pilots 
did exceed limit loads during recovery.
    Because Airbus has chosen to include this optional design feature 
for which part 25 does not contain adequate or appropriate safety 
standards, a proposed special condition pertaining to this feature is 
included. This special condition establishes minimum load factor 
requirements to ensure adequate maneuver capability during normal 
flight.

11. Flight Envelope Protection: High Speed Limiting

    The longitudinal control law design of the A380 incorporates a high 
speed limiting protection system in the normal flight mode. This system 
prevents the pilot from inadvertently or intentionally exceeding the 
airplane maximum design speeds, VD/MD. Part 25 
does not address such a system that would limit or modify flying 
qualities in the high speed region.
    The main features of the high speed limiting function are as 
follows:
     It protects the airplane against high speed/high mach 
number flight conditions beyond VMO/MMO.
     It does not interfere with flight at VMO/
MMO, even in turbulent air.
     It still provides load factor limitation through the 
``pitch limiting'' function described below.
     It restores positive static stability beyond 
VMO/MMO.

[[Page 19019]]

    This special condition establishes requirements to ensure that 
operation of the high speed limiter does not impede normal attainment 
of speeds up to the overspeed warning.

12. Flight Envelope Protection: Pitch and Roll Limiting

    Currently, part 25 does not specifically address flight 
characteristics associated with fixed attitude limits. Airbus proposes 
to implement pitch and roll attitude limiting functions on the A380 via 
the Electronic Flight Control System (EFCS) normal modes. These normal 
modes will prevent airplane pitch attitudes greater than +30 degrees 
and less than -15 degrees and roll angles greater than plus or minus 67 
degrees. In addition, positive spiral stability is introduced for roll 
angles greater than 33 degrees at speeds below VMO/
MMO. At speeds greater than VMO/MMO, 
the maximum aileron control force with positive spiral stability 
results in a maximum bank angle of 45 degrees.
    This special condition establishes requirements to ensure that 
pitch limiting functions do not impede normal maneuvering and that 
pitch and roll limiting functions do not restrict or prevent attaining 
certain roll angles necessary for emergency maneuvering.
    Special conditions to supplement Sec.  25.143 concerning pitch and 
roll limits were developed for the A320, A330 and A340 in which 
performance of the limiting functions was monitored throughout the 
flight test program. The FAA expects similar monitoring to take place 
during the A380 flight test program to substantiate the pitch and roll 
attitude limiting functions and the appropriateness of the chosen 
limits.

13. Flight Envelope Protection: High Incidence Protection and Alpha-
Floor Systems

    The A380 is equipped with a high incidence protection system that 
limits the angle of attack at which the airplane can be flown during 
normal low speed operation and that cannot be overridden by the flight 
crew. The application of this limitation on the angle of attack affects 
the longitudinal handling characteristics of the airplane, so that 
there is no need for the stall warning system during normal operation. 
In addition, the alpha-floor function automatically advances the 
throttles on the operating engines whenever the airplane angle of 
attack reaches a predetermined high value. This function is intended to 
provide increased climb capability. This special condition thus 
addresses the unique features of the low speed high incidence 
protection and the alpha-floor systems on the A380.
    The high incidence protection system prevents the airplane from 
stalling, which means that the stall warning system is not needed 
during normal flight conditions. If there is a failure of the high 
incidence protection system that is not shown to be extremely 
improbable, the flight characteristics at the angle of attack for 
CLMAX must be suitable in the traditional sense, and stall 
warning must be provided in a conventional manner.

14. High Intensity Radiated Fields (HIRF) Protection

    The Airbus Model A380-800 will utilize electrical and electronic 
systems which perform critical functions. These systems may be 
vulnerable to high-intensity radiated fields (HIRF) external to the 
airplane. There is no specific regulation that addresses requirements 
for protection of electrical and electronic systems from HIRF. With the 
trend toward increased power levels from ground-based transmitters and 
the advent of space and satellite communications, coupled with 
electronic command and control of the airplane, the immunity of 
critical avionics/electronics and electrical systems to HIRF must be 
established.
    To ensure that a level of safety is achieved that is equivalent to 
that intended by the regulations incorporated by reference, a special 
condition is needed for the Airbus Model A380 airplane. This special 
condition requires that avionics/electronics and electrical systems 
that perform critical functions be designed and installed to preclude 
component damage and interruption.
    It is not possible to precisely define the HIRF to which the 
airplane will be exposed in service. There is also uncertainty 
concerning the effectiveness of airframe shielding for HIRF. 
Furthermore, coupling of electromagnetic energy to cockpit-installed 
equipment through the cockpit window apertures is undefined. Based on 
surveys and analysis of existing HIRF emitters, adequate protection 
from HIRF exists when there is compliance with either paragraph a. or 
b. below:
    a. A minimum threat of 100 volts rms (root-mean-square) per meter 
electric field strength from 10 KHz to 18 GHz.
    (1) The threat must be applied to the system elements and their 
associated wiring harnesses without the benefit of airframe shielding.
    (2) Demonstration of this level of protection is established 
through system tests and analysis.
    b. A threat external to the airframe of the field strengths 
indicated in the table below for the frequency ranges indicated. Both 
peak and average field strength components from the table below are to 
be demonstrated.

------------------------------------------------------------------------
                                                       Field strength
                                                      (volts per meter)
                     Frequency                     ---------------------
                                                       Peak     Average
------------------------------------------------------------------------
10 kHz-100 kHz....................................         50         50
100 kHz-500 kHz...................................         50         50
500 kHz-2 MHz.....................................         50         50
2 MHz-30 MHz......................................        100        100
30 MHz-70 MHz.....................................         50         50
70 MHz-100 MHz....................................         50         50
100 MHz-200 MHz...................................        100        100
200 MHz-400 MHz...................................        100        100
400 MHz-700 MHz...................................        700         50
700 MHz-1 GHz.....................................        700        100
1 GHz-2 GHz.......................................       2000        200
2 GHz-4 GHz.......................................       3000        200
4 GHz-6 GHz.......................................       3000        200
6 GHz-8 GHz.......................................       1000        200
8 GHz-12 GHz......................................       3000        300
12 GHz-18 GHz.....................................       2000        200
18 GHz-40 GHz.....................................        600       200
------------------------------------------------------------------------
The field strengths are expressed in terms of peak root-mean-square
  (rms) values over the complete modulation period.

    The threat levels identified above are the result of an FAA review 
of existing studies on the subject of HIRF.

15. Operation Without Normal Electrical Power

    These special conditions were developed to address fly-by-wire 
airplanes starting with the Airbus Model A330. As with earlier 
airplanes, the Airbus A380-800 fly-by-wire control system requires a 
continuous source of electrical power for the flight control system to 
remain operable.
    Section 25.1351(d), ``Operation without normal electrical power,'' 
requires safe operation in visual flight rules (VFR) weather conditions 
for at least five minutes with inoperative normal power. This rule was 
structured around a traditional design utilizing mechanical control 
cables for flight control while the crew took time to sort out the 
electrical failure, start the engine(s) if necessary, and re-establish 
some of the electrical power generation capability.
    To maintain the same level of safety as that associated with 
traditional designs, the Model A380 design must not be time limited in 
its operation, including being without the normal source of engine or 
Auxiliary Power Unit (APU) generated electrical power. Service 
experience has shown that the loss of all electrical power generated by 
the airplane's engine generators or APU is not extremely improbable. 
Thus, it must be demonstrated that the airplane

[[Page 19020]]

can continue through safe flight and landing--including steering and 
braking on the ground for airplanes using steer/brake-by-wire--using 
its emergency electrical power systems. These emergency electrical 
power systems must be able to power loads that are essential for 
continued safe flight and landing.

Applicability

    As discussed above, these special conditions are applicable to the 
Airbus A380-800 airplane. Should Airbus apply at a later date for a 
change to the type certificate to include another model incorporating 
the same novel or unusual design features, these special conditions 
would apply to that model as well under the provisions of Sec.  
21.101(a)(1), Amendment 21-69, effective September 16, 1991.

Conclusion

    This action affects only certain novel or unusual design features 
of the Airbus A380-800 airplane. It is not a rule of general 
applicability, and it affects only the applicant which applied to the 
FAA for approval of these features on the airplane.

List of Subjects in 14 CFR Part 25

    Aircraft, Aviation safety, Reporting and recordkeeping 
requirements.

PART 25--[AMENDED]

    The authority citation for these special conditions is as follows:

    Authority: 49 U.S.C. 106(g), 40113, 44701, 44702, 44704.

The Proposed Special Conditions

    Accordingly, pursuant to the authority delegated to me by the 
Administrator, the Federal Aviation Administration (FAA) proposes the 
following special conditions as part of the type certification basis 
for the Airbus A380-800 airplane.

1. Dynamic Braking

    In addition to the requirements of Sec.  25.493(d), the following 
special condition applies:
    Loads arising from the sudden application of maximum braking effort 
must be defined, taking into account the behavior of the braking 
system. Failure conditions of the braking system must be analyzed in 
accordance with the criteria specified in proposed special condition 
number 2, ``Interaction of Systems and Structures.''

2. Interaction of Systems and Structures

    In addition to the requirements of part 25, subparts C and D, the 
following special condition applies:
    a. For airplanes equipped with systems that affect structural 
performance--either directly or as a result of a failure or 
malfunction--the influence of these systems and their failure 
conditions must be taken into account when showing compliance with the 
requirements of part 25, subparts C and D. Paragraph c. below must be 
used to evaluate the structural performance of airplanes equipped with 
these systems.
    b. 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 loads must be treated in accordance with the requirements 
of paragraph a. above.
    c. Interaction of Systems and Structures
    (1) General: The following criteria must be used for showing 
compliance with this special condition and with Sec.  25.629 for 
airplanes equipped with flight control systems, autopilots, stability 
augmentation systems, load alleviation systems, flutter control 
systems, and fuel management systems. If this paragraph is used for 
other systems, it may be necessary to adapt the criteria to the 
specific system.
    (a) The criteria defined herein address only the direct structural 
consequences of the system responses and performances. They cannot be 
considered in isolation but should be included in the overall safety 
evaluation of the airplane. These criteria may, in some instances, 
duplicate standards already established for this evaluation. These 
criteria are applicable only to structures whose failure could prevent 
continued safe flight and landing. Specific criteria that define 
acceptable limits on handling characteristics or stability requirements 
when operating in the system degraded or inoperative modes are not 
provided in this paragraph.
    (b) Depending upon the specific characteristics of the airplane, 
additional studies may be required that go beyond the criteria provided 
in this paragraph in order to demonstrate the capability of the 
airplane to meet other realistic conditions, such as alternative gust 
or maneuver descriptions for an airplane equipped with a load 
alleviation system.
    (c) The following definitions are applicable to this paragraph.
    Structural performance: Capability of the airplane to meet the 
structural requirements of part 25.
    Flight limitations: Limitations that can be applied to the airplane 
flight conditions following an in-flight occurrence and that are 
included in the flight manual (e.g., speed limitations and avoidance of 
severe weather conditions).
    Operational limitations: Limitations, including flight limitations, 
that can be applied to the airplane operating conditions before 
dispatch (e.g., fuel, payload and Master Minimum Equipment List 
limitations).
    Probabilistic terms: The probabilistic terms (probable, improbable, 
and extremely improbable) used in this special condition are the same 
as those used in Sec.  25.1309.
    Failure condition: The term failure condition is the same as that 
used in Sec.  25.1309. However, this special condition applies only to 
system failure conditions that affect the structural performance of the 
airplane (e.g., system failure conditions that induce loads, change the 
response of the airplane to inputs such as gusts or pilot actions, or 
lower flutter margins).
    (2) Effects of Systems on Structures.
    (a) General. The following criteria will be used in determining the 
influence of a system and its failure conditions on the airplane 
structure.
    (b) System fully operative. With the system fully operative, the 
following apply:
    (1) Limit loads must be derived in all normal operating 
configurations of the system from all the limit conditions specified in 
Subpart C, taking into account any special behavior of such a system or 
associated functions or any effect on the structural performance of the 
airplane that may occur up to the limit loads. In particular, any 
significant non-linearity (rate of displacement of control surface, 
thresholds or any other system non-linearities) must be accounted for 
in a realistic or conservative way when deriving limit loads from limit 
conditions.
    (2) The airplane must meet the strength requirements of part 25 
(Static strength, residual strength), using the specified factors to 
derive ultimate loads from the limit loads defined above. The effect of 
non-linearities must be investigated beyond limit conditions to ensure 
that the behavior of the system presents no anomaly compared to the 
behavior below limit conditions. However, conditions beyond limit 
conditions need not be considered, when it can be shown that the 
airplane has design features that will not allow it to exceed those 
limit conditions.
    (3) The airplane must meet the aeroelastic stability requirements 
of Sec.  25.629.
    (c) System in the failure condition. For any system failure 
condition not

[[Page 19021]]

shown to be extremely improbable, the following apply:
    (1) At the time of occurrence. Starting from 1g level flight 
conditions, a realistic scenario, including pilot corrective actions, 
must be established to determine the loads occurring at the time of 
failure and immediately after failure.
    (i) For static strength substantiation, these loads multiplied by 
an appropriate factor of safety that is related to the probability of 
occurrence of the failure are ultimate loads to be considered for 
design. The factor of safety (F.S.) is defined in Figure 1.
[GRAPHIC] [TIFF OMITTED] TP12AP05.005

    (ii) For residual strength substantiation, the airplane must be 
able to withstand two thirds of the ultimate loads defined in 
subparagraph (c)(1)(i) of this section.
    (iii) Freedom from aeroelastic instability must be shown up to the 
speeds defined in Sec.  25.629(b)(2). For failure conditions that 
result in speed increases beyond VC/MC, freedom 
from aeroelastic instability must be shown to increased speeds, so that 
the margins intended by Sec.  25.629(b)(2) are maintained.
    (iv) Failures of the system that result in forced structural 
vibrations (oscillatory failures) must not produce loads that could 
result in detrimental deformation of primary structure.
    (2) For the continuation of the flight. For the airplane in the 
system failed state and considering any appropriate reconfiguration and 
flight limitations, the following apply:
    (i) The loads derived from the following conditions at speeds up to 
VC or the speed limitation prescribed for the remainder of 
the flight must be determined:
    (A) the limit symmetrical maneuvering conditions specified in Sec.  
25.331 and in Sec.  25.345.
    (B) the limit gust and turbulence conditions specified in Sec.  
25.341 and in Sec.  25.345.
    (C) the limit rolling conditions specified in Sec.  25.349 and the 
limit unsymmetrical conditions specified in Sec.  25.367 and Sec.  
25.427(b) and (c).
    (D) the limit yaw maneuvering conditions specified in Sec.  25.351.
    (E) the limit ground loading conditions specified in Sec.  25.473 
and Sec.  25.491.
    (ii) For static strength substantiation, each part of the structure 
must be able to withstand the loads in subparagraph (2)(i) of this 
paragraph multiplied by a factor of safety, depending on the 
probability of being in this failure state. The factor of safety is 
defined in Figure 2.

[[Page 19022]]

[GRAPHIC] [TIFF OMITTED] TP12AP05.006

Q j = (Tj)(Pj)
Where:

Tj = Average time spent in failure condition j (in hours)
Pj = Probability of occurrence of failure mode j (per hour)


    Note: If Pj is greater than 10-\3\ per 
flight hour, then a 1.5 factor of safety must be applied to all 
limit load conditions specified in Subpart C.


    (iii) For residual strength substantiation, the airplane must be 
able to withstand two thirds of the ultimate loads defined in 
subparagraph (c)(2)(ii).
    (iv) If the loads induced by the failure condition have a 
significant effect on fatigue or damage tolerance, then their effects 
must be taken into account.
    (v) Freedom from aeroelastic instability must be shown up to a 
speed determined from Figure 3. Flutter clearance speeds V' and V'' may 
be based on the speed limitation specified for the remainder of the 
flight, using the margins defined by Sec.  25.629(b).
[GRAPHIC] [TIFF OMITTED] TP12AP05.007

V' = Clearance speed as defined by Sec.  25.629(b)(2).
V'' = Clearance speed as defined by Sec.  25.629(b)(1).
Q j = (Tj)(Pj)
Where:
Tj = Average time spent in failure condition j (in hours)
Pj = Probability of occurrence of failure mode j (per hour)

    Note: If Pj is greater than 10-\3\ per 
flight hour, then the flutter clearance speed must not be less than 
V''.


    (vi) Freedom from aeroelastic instability must also be shown up to 
V' in Figure 3 above for any probable system failure condition combined 
with any damage required or selected for investigation by Sec.  
25.571(b).
    (3) Consideration of certain failure conditions may be required by 
other sections of this Part, regardless of calculated system 
reliability. Where analysis shows the probability of these failure 
conditions to be less than 10-\9\, criteria other than those 
specified in this paragraph may be used for structural substantiation 
to show continued safe flight and landing.
    (d) Warning considerations. For system failure detection and 
warning, the following apply:
    (1) The system must be checked for failure conditions, not 
extremely improbable, that degrade the structural capability below the 
level required by part 25 or significantly reduce the reliability of 
the remaining system. The flight crew must be made aware of these 
failures before flight. Certain elements of the control system, such as

[[Page 19023]]

mechanical and hydraulic components, may use special periodic 
inspections, and electronic components may use daily checks in lieu of 
warning systems to achieve the objective of this requirement. These 
certification maintenance requirements must be limited to components 
that are not readily detectable by normal warning systems and where 
service history shows that inspections will provide an adequate level 
of safety.
    (2) The existence of any failure condition, not extremely 
improbable, during flight that could significantly affect the 
structural capability of the airplane and for which the associated 
reduction in airworthiness can be minimized by suitable flight 
limitations must be signaled to the flightcrew. For example, failure 
conditions that result in a factor of safety between the airplane 
strength and the loads of part 25, subpart C below 1.25 or flutter 
margins below V'' must be signaled to the crew during flight.
    (e) Dispatch with known failure conditions. If the airplane is to 
be dispatched in a known system failure condition that affects 
structural performance or affects the reliability of the remaining 
system to maintain structural performance, then the provisions of this 
special condition must be met for the dispatched condition and for 
subsequent failures. Flight limitations and expected operational 
limitations may be taken into account in establishing Qj as the 
combined probability of being in the dispatched failure condition and 
the subsequent failure condition for the safety margins in Figures 2 
and 3. These limitations must be such that the probability of being in 
this combined failure state and then subsequently encountering limit 
load conditions is extremely improbable. No reduction in these safety 
margins is allowed, if the subsequent system failure rate is greater 
than 1E-3 per flight hour.

3. Limit Pilot Forces

    In addition to the requirements of Sec.  25.397(c) the following 
special condition applies:
    The limit pilot forces are as follows:
    a. For all components between and including the handle and its 
control stops.

------------------------------------------------------------------------
                   Pitch                                Roll
------------------------------------------------------------------------
Nose up 200 lbf...........................  Nose left 100 lbf.
Nose down 200 lbf.........................  Nose right 100 lbf.
------------------------------------------------------------------------

    b. For all other components of the side stick control assembly, but 
excluding the internal components of the electrical sensor assemblies 
to avoid damage as a result of an in-flight jam.

------------------------------------------------------------------------
                   Pitch                                Roll
------------------------------------------------------------------------
Nose up 125 lbf...........................  Nose left 50 lbf.
Nose down 125 lbf.........................  Nose right 50 lbf.
------------------------------------------------------------------------

4. Side Stick Controllers

    In the absence of specific requirements for side stick controllers, 
the following special condition applies:
    a. Pilot strength: In lieu of the ``strength of pilots'' limits 
shown in Sec.  25.143(c) for pitch and roll and in lieu of the specific 
pitch force requirements of Sec. Sec.  25.145(b) and 25.175(d), it must 
be shown that the temporary and maximum prolonged force levels for the 
side stick controllers are suitable for all expected operating 
conditions and configurations, whether normal or non-normal.
    b. Pilot control authority: The electronic side stick controller 
coupling design must provide for corrective and/or overriding control 
inputs by either pilot with no unsafe characteristics. Annunciation of 
the controller status must be provided and must not be confusing to the 
flight crew.
    c. Pilot control: It must be shown by flight tests that the use of 
side stick controllers does not produce unsuitable pilot-in-the-loop 
control characteristics when considering precision path control/ tasks 
and turbulence. In addition, pitch and roll control force and 
displacement sensitivity must be compatible, so that normal inputs on 
one control axis will not cause significant unintentional inputs on the 
other.
    d. Autopilot quick-release control location: In lieu of compliance 
with 25.1329(d), autopilot quick release (emergency) controls must be 
on both side stick controllers. The quick release means must be located 
so that it can readily and easily be used by the flight crew.

5. Dive Speed Definition

    In lieu of the requirements of Sec.  25.335(b)(1)--if the flight 
control system includes functions which act automatically to initiate 
recovery before the end of the 20 second period specified in Sec.  
25.335(b)(1)--the greater of the speeds resulting from the following 
special condition applies.
    a. From an initial condition of stabilized flight at VC/
MC, the airplane is upset so as to take up a new flight path 
7.5 degrees below the initial path. Control application, up to full 
authority, is made to maintain this new flight path. Twenty seconds 
after initiating the upset, manual recovery is made at a load factor of 
1.5 g (0.5 acceleration increment) or such greater load factor that is 
automatically applied by the system with the pilot's pitch control 
neutral. 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 
recovery is made, at which time power reduction and the use of pilot 
controlled drag devices may be used.
    b. From a speed below VC/MC with power to 
maintain stabilized level flight at this speed, the airplane is upset 
so as to accelerate through VC/MC at a flight 
path 15 degrees below the initial path--or at the steepest nose down 
attitude that the system will permit with full control authority if 
less than 15 degrees.


    Note: The pilot's controls may be in the neutral position after 
reaching VC/MC and before recovery is 
initiated.

    c. Recovery may be initiated three seconds after operation of high 
speed warning system by application of a load of 1.5g (0.5 acceleration 
increment) or such greater load factor that is automatically applied by 
the system with the pilot's pitch control neutral. Power may be reduced 
simultaneously. All other means of decelerating the airplane, the use 
of which is authorized up to the highest speed reached in the maneuver, 
may be used. The interval between successive pilot actions must not be 
less than one second.
    d. The applicant must also demonstrate either that
    (1) the speed margin, established as above, will not be exceeded in 
inadvertent or gust induced upsets, resulting in initiation of the dive 
from non-symmetric attitudes, or
    (2) the airplane is protected by the flight control laws from 
getting into non-symmetric upset conditions.
    e. The probability of failure of the protective system that 
mitigates for the reduced speed margin must be less than 
10-5 per flight hour, except that the probability of failure 
may be greater than 10-5, but not greater than 
10-3, per flight hour, provided that:
    (1) Failures of the system are annunciated to the pilots, and
    (2) The flight manual instructions require the pilots to reduce the 
speed of the airplane to a value that maintains a speed margin between 
VMO and VD consistent with showing compliance 
with 25.335(b) without the benefit of the system, and
    (3) no dispatch of the airplane is allowed with the system 
inoperative.

[[Page 19024]]

6. Electronic Flight Control System: Lateral-Directional and 
Longitudinal Stability and Low Energy Awareness

    In lieu of the requirements of Sec.  25.171 and sub-section 
25.177(c), the following special condition applies:
    a. The airplane must be shown to have suitable static lateral, 
directional, and longitudinal stability in any condition normally 
encountered in service, including the effects of atmospheric 
disturbance.
    b. The airplane must provide adequate awareness to the pilot of a 
low energy (low speed/low thrust/low height) state when fitted with 
flight control laws presenting neutral longitudinal stability 
significantly below the normal operating speeds.
    c. 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 symmetrical power condition, at 
speeds from 1.13 VS1g up to VFE, VLE, 
or VFC/MFC (as appropriate).
    d. In straight, steady sideslips (unaccelerated forward slips), the 
rudder control movements and forces must be substantially proportional 
to the angle of sideslip, and the factor of proportionality must be 
between limits found necessary for safe operation throughout the range 
of sideslip angles appropriate to the operation of the airplane. At 
greater angles--up to the angle at which full rudder control is used or 
a rudder pedal force of 180 pounds (81.72 kg) is obtained--the rudder 
pedal forces may not reverse, and increased rudder deflection must 
produce increased angles of sideslip. Unless the airplane has a 
suitable sideslip indication, there must be enough bank and lateral 
control deflection and force accompanying sideslipping to clearly 
indicate any departure from steady, unyawed flight.

7. Electronic Flight Control System: Control Surface Awareness

    In addition to the requirements of Sec. Sec.  25.143, 25.671 and 
25.672, the following special condition applies:
    a. A suitable flight control position annunciation must be provided 
to the crew in the following situation:
    A flight condition exists in which--without being commanded by the 
crew--control surfaces are coming so close to their limits that return 
to normal flight and (or) continuation of safe flight requires a 
specific crew action.
    b. In lieu of control position annunciation, existing indications 
to the crew may be used to prompt crew action, if they are found to be 
adequate.


    Note: The term ``suitable'' also indicates an appropriate 
balance between nuisance and necessary operation.


8. Electronic Flight Control System: Flight Characteristics Compliance 
Via the Handling Quantities Rating Method (HQRM)

    a. Flight Characteristics Compliance Determination for EFCS Failure 
Cases:
    In lieu of compliance with Sec.  25.672(c), the HQRM contained in 
Appendix 7 of AC 25-7A must be used for evaluation of EFCS 
configurations resulting from single and multiple failures not shown to 
be extremely improbable.
    The handling qualities ratings are as follows:
    (1) Satisfactory: Full performance criteria can be met with routine 
pilot effort and attention.
    (2) Adequate: Adequate for continued safe flight and landing; full 
or specified reduced performance can be met, but with heightened pilot 
effort and attention.
    (3) Controllable: Inadequate for continued safe flight and landing, 
but controllable for return to a safe flight condition, safe flight 
envelope and/or reconfiguration, so that the handling qualities are at 
least Adequate.
    b. Handling qualities will be allowed to progressively degrade with 
failure state, atmospheric disturbance level, and flight envelope, as 
shown in Figure 12 of Appendix 7. Specifically, for probable failure 
conditions within the normal flight envelope, the pilot-rated handling 
qualities must be satisfactory in light atmospheric disturbance and 
adequate in moderate atmospheric disturbance. The handling qualities 
rating must not be less than adequate in light atmospheric disturbance 
for improbable failures.


    Note: AC 25-7A, Appendix 7 presents a method of compliance and 
provides guidance for the following:
     Minimum handling qualities rating requirements in 
conjunction with atmospheric disturbance levels, flight envelopes, 
and failure conditions (Figure 12),
     Flight Envelope definition (Figures 5A, 6 and 7),
     Atmospheric Disturbance Levels (Figure 5B),
     Flight Control System Failure State (Figure 5C),
     Combination Guidelines (Figures 5D, 9 and 10), and
     General flight task list, from which appropriate 
specific tasks can be selected or developed (Figure 11).

9. Flight Envelope Protection

    a. General Limiting Requirements: (1) Onset characteristics of each 
envelope protection feature must be smooth, appropriate to the phase of 
flight and type of maneuver, and not in conflict with the ability of 
the pilot to satisfactorily change the airplane flight path, speed, or 
attitude, as needed.
    (2) Limit values of protected flight parameters (and if applicable, 
associated warning thresholds) must be compatible with the following:
    (a) Airplane structural limits,
    (b) Required safe and controllable maneuvering of the airplane, and
    (c) Margins to critical conditions. Dynamic maneuvering, airframe 
and system tolerances (both manufacturing and in-service), and non-
steady atmospheric conditions--in any appropriate combination and phase 
of flight--must not result in a limited flight parameter beyond the 
nominal design limit value that would cause unsafe flight 
characteristics.
    (3) The airplane must be responsive to intentional dynamic 
maneuvering to within a suitable range of the parameter limit. Dynamic 
characteristics, such as damping and overshoot, must also be 
appropriate for the flight maneuver and limit parameter in question.
    (4) When simultaneous envelope limiting is engaged, adverse 
coupling or adverse priority must not result.
    b. Failure States: EFCS failures, including sensor failures, must 
not result in a condition where a parameter is limited to such a 
reduced value that safe and controllable maneuvering is no longer 
available. The crew must be alerted by suitable means, if any change in 
envelope limiting or maneuverability is produced by single or multiple 
failures of the EFCS not shown to be extremely improbable.

10. Flight Envelope Protection: Normal Load Factor (g) Limiting

    In addition to the requirements of 25.143(a)--and in the absence of 
other limiting factors--the following special condition applies:
    a. The positive limiting load factor must not be less than:
    (1) 2.5g for the EFCS normal state.
    (2) 2.0g for the EFCS normal state with the high lift devices 
extended.
    b. The negative limiting load factor must be equal to or more 
negative than:
    (1) Minus 1.0g for the EFCS normal state.
    (2) 0.0g for the EFCS normal state with high lift devices extended.

    Note: This Special Condition does not impose an upper bound for 
the normal load factor limit, nor does it require that the limit 
exist. If the limit is set at a value beyond the structural design 
limit maneuvering load factor ``n,'' indicated in Sec.  25.333(b) 
and 25.337(b) and (c), there should be a very positive tactile feel 
built into the controller

[[Page 19025]]

and obvious to the pilot that serves as a deterrent to inadvertently 
exceeding the structural limit.

11. Flight Envelope Protection High Speed Limiting

    In addition to Sec.  25.143, the following special condition 
applies:
    Operation of the high speed limiter during all routine and descent 
procedure flight must not impede normal attainment of speeds up to the 
overspeed warning.

12. Flight Envelope Protection: Pitch and Roll Limiting

    In addition to Sec.  25.143, the following special condition 
applies:
    a. The pitch limiting function must not impede normal maneuvering 
for pitch angles up to the maximum required for normal maneuvering--
including a normal all-engines operating takeoff plus a suitable margin 
to allow for satisfactory speed control.
    b. The pitch and roll limiting functions must not restrict or 
prevent attaining roll angles up to 65 degrees or pitch attitudes 
necessary for emergency maneuvering. Spiral stability, which is 
introduced above 33 degrees roll angle, must not require excessive 
pilot strength to achieve roll angles up to 65 degrees.

13. Flight Envelope Protection: High Incidence Protection and Alpha-
Floor Systems

    a. Definitions. For the purpose of this special condition, the 
following definitions apply:
    High Incidence Protection System. A system that operates directly 
and automatically on the airplane's flying controls to limit the 
maximum angle of attack that can be attained to a value below that at 
which an aerodynamic stall would occur.
    Alpha-Floor System. A system that automatically increases thrust on 
the operating engines when the angle of attack increases through a 
particular value.
    Alpha Limit. The maximum angle of attack at which the airplane 
stabilizes with the high incidence protection system operating and the 
longitudinal control held on its aft stop.
    Vmin The minimum steady flight speed is the stabilized, calibrated 
airspeed obtained when the airplane is decelerated at an entry rate not 
exceeding 1 knot per second, until the longitudinal pilot control is on 
its stop with the high incidence protection system operating.
    Vmin1g Vmin corrected to 1g conditions. It is the minimum 
calibrated airspeed at which the airplane can develop a lift force 
normal to the flight path and equal to its weight when at an angle of 
attack not greater than that determined for Vmin.
    b. Capability and Reliability of the High Incidence Protection 
System:
    (1) It must not be possible to encounter a stall during pilot 
induced maneuvers, and handling characteristics must be acceptable, as 
required by Paragraphs e and f below, entitled High Incidence Handling 
Demonstrations and High Incidence Handling Characteristics 
respectively.
    (2) The airplane must be protected against stalling due to the 
effects of windshears and gusts at low speeds, as required by Paragraph 
g below, entitled Atmospheric Disturbances.
    (3) The ability of the high incidence protection system to 
accommodate any reduction in stalling incidence resulting from residual 
ice must be verified.
    (4) The reliability of the system and the effects of failures must 
be acceptable, in accordance with Sec.  25.1309 and Advisory Circular 
25.1309-1A, System Design and Analysis.
    (5) The high incidence protection system must not impede normal 
maneuvering for pitch angles up to the maximum required for normal 
maneuvering, including a normal all-engines operating takeoff plus a 
suitable margin to allow for satisfactory speed control.
    c. Minimum Steady Flight Speed and Reference Stall Speed:
    In lieu of the requirements of Sec.  25.103, the following special 
condition applies:
    (1) Vmin The minimum steady flight speed, for the 
airplane configuration under consideration and with the high incidence 
protection system operating, is the final stabilized calibrated 
airspeed obtained when the airplane is decelerated at an entry rate not 
exceeding 1 knot per second until the longitudinal pilot control is on 
its stop.
    (2) The minimum steady flight speed, Vmin, must be 
determined with:
    (a) The high incidence protection system operating normally.
    (b) Idle thrust.
    (c) Alpha-floor system inhibited.
    (d) All combinations of flap settings and landing gear positions.
    (e) The weight used when VSR is being used as a factor 
to determine compliance with a required performance standard.
    (f) The most unfavorable center of gravity allowable, and
    (g) The airplane trimmed for straight flight at a speed achievable 
by the automatic trim system.
    (3) Vmin1g is Vmin corrected to 1g 
conditions. Vmin1g is the minimum calibrated airspeed at 
which the airplane can develop a lift force normal to the flight path 
and equal to its weight when at an angle of attack not greater than 
that determined for Vmin. Vmin1g is defined as 
follows:
[GRAPHIC] [TIFF OMITTED] TP12AP05.008

Where:
n z w = load factor normal to the flight path at 
Vmin

    (4) The Reference Stall Speed, VSR, is a calibrated 
airspeed selected by the applicant. VSR may not be less than 
the 1g stall speed. VSR is expressed as:
[GRAPHIC] [TIFF OMITTED] TP12AP05.009

Where:
VCLMAX = Calibrated airspeed obtained when the load factor-
corrected lift coefficient
[GRAPHIC] [TIFF OMITTED] TP12AP05.010

is first a maximum during the maneuver prescribed in paragraph (5)(h) 
of this section.

nzw = Load factor normal to the flight path at 
VCLMAX
W = Airplane gross weight
S = Aerodynamic reference wing area, and
q = Dynamic pressure.

    (5) VCLMAX must be determined with the following 
conditions:
    (a) Engines idling or--if that resultant thrust causes an 
appreciable decrease in stall speed--not more than zero thrust at the 
stall speed
    (b) The airplane in other respects, such as flaps and landing gear, 
in the condition existing in the test or performance standard in which 
VSR is being used.
    (c) The weight used when VSR is being used as a factor 
to determine compliance with a required performance standard.
    (d) The center of gravity position that results in the highest 
value of reference stall speed.
    (e) The airplane trimmed for straight flight at a speed achievable 
by the automatic trim system, but not less than 1.13 VSR and 
not greater than 1.3 VSR.
    (f) The alpha-floor system inhibited.
    (g) The high incidence protection system adjusted to a high enough 
incidence to allow full development of the 1g stall.
    (h) Starting from the stabilized trim condition, apply the 
longitudinal control to decelerate the airplane so that

[[Page 19026]]

the speed reduction does not exceed one knot per second.
    (6) The flight characteristics at the angle of attack for 
CLMAX must be suitable in the traditional sense at FWD and 
AFT CG in straight and turning flight at IDLE power. Although for a 
normal production EFCS and steady full aft stick this angle of attack 
for CLMAX cannot be achieved, the angle of attack can be 
obtained momentarily under dynamic circumstances and deliberately in a 
steady state sense with some EFCS failure conditions.
    d. Stall Warning. (1) Normal Operation: If the conditions of 
Paragraph b above which is entitled Capability and Reliability of the 
High Incidence Protection System are satisfied, a level of safety 
equivalent to that intended by Sec.  25.207, Stall Warning, must be 
considered to have been met without provision of an additional, unique 
warning device.
    (2) Failure Cases: Following failures of the high incidence 
protection system not shown to be extremely improbable, if the system 
no longer satisfies sub paragraphs (1), (2), and (3) of Paragraph b 
above which is entitled Capability and Reliability of the High 
Incidence Protection System, stall warning must be provided in 
accordance with Sec.  25.207. The stall warning should prevent 
inadvertent stall under the following conditions:
    (a) Power off straight stall approaches to a speed 5 percent below 
the warning onset.
    (b) Turning flight stall approaches at entry rates up to 3 knots 
per second when recovery is initiated not less than one second after 
the warning onset.
    e. High Incidence Handling Demonstrations: In lieu of the 
requirements of Sec.  25.201, the following special condition applies:
    Maneuvers to the limit of the longitudinal control in the nose up 
direction must be demonstrated in straight flight and in 30 degree 
banked turns under the following conditions:
    (1) The high incidence protection system operating normally.
    (2) Initial power condition of:
    (a) Power off
    (b) The power necessary to maintain level flight at 1.5 
VSR1, where VSR1 is the reference stall speed 
with the flaps in the approach position, the landing gear retracted, 
and the maximum landing weight. The flap position to be used to 
determine this power setting is that position in which the stall speed, 
VSR1, does not exceed 110% of the stall speed, 
VSR0, with the flaps in the most extended landing position.
    (3) Alpha-floor system operating normally, unless more severe 
conditions are achieved with alpha-floor inhibited.
    (4) Flaps, landing gear and deceleration devices in any likely 
combination of positions.
    (5) Representative weights within the range for which certification 
is requested, and
    (6) The airplane trimmed for straight flight at a speed achievable 
by the automatic trim system.
    (7) Starting at a speed sufficiently above the minimum steady 
flight 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 control reaches the stop.
    (8) The longitudinal control must be maintained at the stop until 
the airplane has reached a stabilized flight condition and must then be 
recovered by normal recovery techniques.
    (9) The requirements for turning flight maneuver demonstrations 
must also be met with accelerated rates of entry to the incidence 
limit, up to the maximum rate achievable.
    f. High Incidence Handling Characteristics: In lieu of the 
requirements of Sec.  25.203, the following special condition applies:
    (1) Throughout maneuvers with a rate of deceleration of not more 
than 1 knot per second, both in straight flight and in 30 degree banked 
turns, the airplane's characteristics must be as follows:
    (a) There must not be any abnormal airplane nose-up pitching.
    (b) There must not be any uncommanded nose-down pitching that would 
be indicative of stall. However, reasonable attitude changes associated 
with stabilizing the incidence at alpha limit as the longitudinal 
control reaches the stop would be acceptable. Any reduction of pitch 
attitude associated with stabilizing the incidence at the alpha limit 
should be achieved smoothly and at a low pitch rate, such that it is 
not likely to be mistaken for natural stall identification.
    (c) There must not be any uncommanded lateral or directional 
motion, and the pilot must retain good lateral and directional control 
by conventional use of the cockpit controllers throughout the maneuver.
    (d) The airplane must not exhibit buffeting of a magnitude and 
severity that would act as a deterrent to completing the maneuver.
    (2) In maneuvers with increased rates of deceleration, some 
degradation of characteristics is acceptable, associated with a 
transient excursion beyond the stabilized alpha-limit. However, the 
airplane must not exhibit dangerous characteristics or characteristics 
that would deter the pilot from holding the longitudinal controller on 
the stop for a period of time appropriate to the maneuvers.
    (3) It must always be possible to reduce incidence by conventional 
use of the controller.
    (4) The rate at which the airplane can be maneuvered from trim 
speeds associated with scheduled operating speeds such as V2 
and VREF up to alpha-limit must not be unduly damped or 
significantly slower than can be achieved on conventionally controlled 
transport airplanes.
    g. Atmospheric Disturbances: Operation of the high incidence 
protection system and the alpha-floor system must not adversely affect 
aircraft control during expected levels of atmospheric disturbances or 
impede the application of recovery procedures in case of windshear. 
Simulator tests and analysis may be used to evaluate such conditions 
but must be validated by limited flight testing to confirm handling 
qualities at critical loading conditions.
    h. Alpha Floor: The alpha-floor setting must be such that the 
aircraft can be flown at normal landing operational speed and 
maneuvered up to bank angles consistent with the flight phase, 
including the maneuver capabilities specified in 25.143(g), without 
triggering alpha-floor. In addition, there must be no alpha-floor 
triggering, unless appropriate, when the airplane is flown in usual 
operational maneuvers and in turbulence.
    i. Proof of Compliance: In addition to the requirements of Sec.  
25.21, the following special condition applies:
    The flying qualities must be evaluated at the most unfavorable 
center of gravity position.
    j. Longitudinal Control: (1) In lieu of the requirements of Sec.  
25.145(a) and 25.145(a)(1), the following special condition applies:
    It must be possible--at any point between the trim speed for 
straight flight achievable by the automatic trim system and 
Vmin--to pitch the nose downward, so that the acceleration 
to this selected trim speed is prompt, with the airplane trimmed for 
straight flight at the speed achievable by the automatic trim system.
    (2) In lieu of the requirements of Sec.  25.145(b)(6), the 
following special condition applies:
    With power off, flaps extended and the airplane trimmed at 1.3 
VSR1, obtain and maintain airspeeds between Vmin 
and either 1.6 VSR1 or VFE, whichever is lower.

[[Page 19027]]

    k. Airspeed Indicating System: (1) In lieu of the requirements of 
subsection 25.1323(c)(1), the following special condition applies:
    VMO to Vmin with the flaps retracted.
    (2) In lieu of the requirements of subsection 25.1323(c)(2), the 
following special condition applies:
    Vmin to VFE with flaps in the landing 
position.

14. High Intensity Radiated Fields (HIRF) Protection

    a. Protection from Unwanted Effects of High-intensity Radiated 
Fields:
    Each electrical and electronic system which performs critical 
functions must be designed and installed to ensure that the operation 
and operational capabilities of these systems to perform critical 
functions are not adversely affected when the airplane is exposed to 
high intensity radiated fields external to the airplane.
    b. For the purposes of this special condition, the following 
definition applies:
    Critical Functions: Functions whose failure would contribute to or 
cause a failure condition which would prevent the continued safe flight 
and landing of the airplane.

15. Operation Without Normal Electrical Power

    In lieu of the requirements of Sec.  25.1351(d), the following 
special condition applies:
    It must be demonstrated by test or combination of test and analysis 
that the airplane can continue safe flight and landing with inoperative 
normal engine and APU generator electrical power (i.e., electrical 
power sources, excluding the battery and any other standby electrical 
sources). The airplane operation should be considered at the critical 
phase of flight and include the ability to restart the engines and 
maintain flight for the maximum diversion time capability being 
certified.

    Issued in Renton, Washington, on March 29, 2005.
Kalene C. Yanamura,
Acting Manager, Transport Airplane Directorate, Aircraft Certification 
Service.
[FR Doc. 05-7320 Filed 4-11-05; 8:45 am]
BILLING CODE 4910-13-P