[Federal Register Volume 65, Number 157 (Monday, August 14, 2000)]
[Proposed Rules]
[Pages 49513-49522]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 00-20584]


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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 23

[Docket No. CE162; Notice No. 23-00-03-SC]


Special Conditions: Ayres Corporation, Model LM 200, 
``Loadmaster''; Propulsion

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Notice of proposed special conditions.

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SUMMARY: This document proposes special conditions for the Ayres 
Corporation, Model LM 200 airplane. This airplane will have a novel or 
unusual design feature associated with a 14 CFR part 23 commuter 
category airplane incorporating a propulsion system that consists of 
two turboshaft engines driving a single propeller through a combining 
gearbox. The

[[Page 49514]]

applicable airworthiness regulations do not contain adequate or 
appropriate safety standards for this design feature. These proposed 
special conditions 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.

DATES: Comments must be received on or before September 13, 2000.

ADDRESSES: Comments on this proposal may be mailed in duplicate to: 
Federal Aviation Administration, Regional Counsel, ACE-7, Attention: 
Rules Docket, Docket No. CE162, 901 Locust St., Kansas City, Missouri 
64106, or delivered in duplicate to the Regional Counsel at the above 
address. Comments must be marked: CE162. 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: Mr. Brian Hancock, Federal Aviation 
Administration, Aircraft Certification Service, Small Airplane 
Directorate, ACE-112, 901 Locust Street, Kansas City, Missouri, 816-
329-4143, fax 816-329-4090.

SUPPLEMENTARY INFORMATION:

Comments Invited

    Interested persons are invited to participate in the making of 
these proposed special conditions by submitting such written data, 
views, or arguments as they may desire. Communications should identify 
the regulatory docket or notice number and be submitted in duplicate to 
the address specified above. All communications received on or before 
the closing date for comments will be considered by the Administrator. 
The proposals described in this notice may be changed in light of the 
comments received. All comments received will be available in the Rules 
Docket for examination by interested persons, both before and after the 
closing date for comments. A report summarizing each substantive public 
contact with FAA personnel concerning this rulemaking will be filed in 
the docket. Persons wishing the FAA to acknowledge receipt of their 
comments submitted in response to this notice must include with those 
comments a self-addressed, stamped postcard on which the following 
statement is made: ``Comments to Docket No. CE162.'' The postcard will 
be date stamped and returned to the commenter.

Background

    On April 16, 1996, Ayres Corporation applied for a type certificate 
for their new Model LM 200 and reapplied in May 1997 adding passenger 
and combi configurations. The Model LM 200 airplane will have a 19,000 
pound maximum takeoff weight with a payload capacity about 7,500 
pounds. The propulsion system will consist of an LHTEC CTP800-4T 
powerplant driving a single Hamilton Standard Model 568F-11, 12.9-foot 
diameter, propeller. The powerplant consists of two LHTEC CTS800 
derivative turboshaft engines plus a combining gearbox. The powerplant 
will be certified to 14 CFR part 33 identified as a twin power section 
turboshaft assembly. The two turboshaft engines will be certified as 
part of the twin power section turboshaft assembly (powerplant) and 
will not have separate individual type certificates. The airplane will 
be of conventional, semi-monocoque, aluminum construction with a high 
cantilever wing, fixed gear, mechanical and electro-mechanical controls 
and will be unpressurized. Certification will include flight into known 
icing and single pilot, IFR operations. Three interior configurations 
have been proposed: a cargo configuration (bulk or containerized 
cargo), a nine-passenger configuration, and ``combi'' (combination of 
passenger and cargo).

Type Certification Basis

    Under the provisions of 14 CFR part 21, Sec. 21.17, Ayres 
Corporation must show that the Model LM 200 meets the applicable 
provisions of part 23 as amended by Amendments 23-1 through Amendment 
53, effective April 30, 1998.
    If the Administrator finds that the applicable airworthiness 
regulations (i.e., part 23) do not contain adequate or appropriate 
safety standards for the Model LM 200 because of a novel or unusual 
design feature, special conditions are prescribed under the provisions 
of Sec. 21.16.
    In addition to the applicable airworthiness regulations and special 
conditions, the Model LM 200 must comply with the part 23 fuel vent and 
exhaust emission requirements of 14 CFR part 34, the part 23 noise 
certification requirements of 14 CFR part 36, and the FAA must issue a 
finding of regulatory adequacy pursuant to Sec. 611 of Public Law 92-
574, the ``Noise Control Act of 1972.''
    Special conditions, as appropriate, are issued in accordance with 
Sec. 11.49 after public notice, as required by Secs. 11.28 and 
11.29(b), and become part of the type certification basis in accordance 
with Sec. 21.17(a)(2).
    Special conditions are initially applicable to the model for which 
they are issued. Should the type certificate for the Model LM 200 be 
amended later to include any other model that incorporates the same 
novel or unusual design feature, these special conditions would also 
apply to the other model under the provisions of Sec. 21.101(a)(1).

Novel or Unusual Design Features

    The following definitions will apply to the Ayres Model LM 200 
airplane design:
    Powerplant--The LHTEC model CTP800-4T powerplant, consists of two 
CTS800 derivative turboshaft engines, a GKN Westland combining gearbox 
(CGB), and the engine assembly support structure. The powerplant is 
capable of providing 2,700 shp combined output power at takeoff and 
1,350 shp with one engine inoperative. The CTP800-4T powerplant will 
obtain a part 33 type certificate identifying the powerplant as a 
``twin power section turboshaft assembly.''
    Engine--An LHTEC CTS800 derivative, non-regenerative, front drive, 
free turbine power section, which includes compressor, combustor, 
turbine and accessories group. Each engine of the CTP800-4T is 
separately controlled by a fully redundant full authority digital 
electronic control (FADEC). The two engines will only be certified as 
part of the CTP800-4T powerplant. The CTP800-4T type certificate data 
sheet will include ratings and limitations for each engine in addition 
to that of the powerplant.
    Engine Assembly Support Structure--The supporting structure that 
connects the two engines to the CGB. This structure will be type 
certificated as part of the CTP800-4T powerplant under part 33.
    Propulsion System Unit (PSU)--The Model LM 200 airplane PSU 
consists of the powerplant plus the airframe mounted non-integrated 
lubrication system components, which include the CGB oil tank and CGB/
engine oil cooler, as well as a single Hamilton Sundstrand Model 568F-
11 propeller system.
    Combining Gearbox (CGB)--All components necessary to transmit power 
from the two engines to the propeller. This includes couplings, 
supporting bearings for shafts, brake assemblies, clutches, gearboxes, 
transmissions, any attached accessory pads or drives, and any cooling 
fans that are attached to, or mounted on, the CGB. The CGB will be type 
certificated as part of the CTP800-4T powerplant under part 33.
    Multi-Engine--For the Model LM 200 and its powerplant 
configuration, ``multi-engine'' refers to the twin engine capability 
and ratings of the CTP800-4T powerplant in regard to type

[[Page 49515]]

certification in the commuter category and flight operation.
    One Engine Inoperative (OEI)--For the LM 200 airplane, ``one engine 
inoperative'' refers to a condition in which one engine of the CTP800-
4T powerplant is not operational and the operation of the propeller is 
unchanged.
    Part 23 does not contain adequate or appropriate requirements for 
the Ayres Model LM 200 powerplant installation of twin engines driving 
a single propeller through a combining gearbox. Issues include 
preventing unbalance damage to either the engines or the powerplant 
mounting system, or both, resulting from any engine or propeller single 
failure or probable combination of failures and the capability to 
continue safe flight to a landing. The propeller and other non-
redundant components must be of sufficient durability to minimize any 
possibility of a failure that could have catastrophic implications to 
either the airplane or its propulsion system, or both.
    Elements of these proposed special conditions have been developed 
to supplement part 23 standards that are considered inadequate to 
address the Model LM 200 airplane design, namely Secs. 23.53, 23.67, 
23.69, 23.75, 23.77, 23.903, 23.1191, 23.1305, 23.1583, 23.1585 and 
23.1587.
    Special Conditions addressing the engine isolation requirements of 
Sec. 23.903 were not included, as the current rule is considered 
adequate. However, since the design of the multi-engine, single 
propeller Model LM 200 airplane will be significantly affected by this 
rule, the following comments are provided. Section 23.903(c) states, 
``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 crew member 
for continued safe operation of the remaining engines.'' This is a 
fail-safe requirement in that it takes advantage of the redundancy 
provided by having multiple engines that are physically separated from 
each other, which is intended to ensure that no single failure 
affecting one engine will result in the loss of the airplane (also 
reference Sec. 23.903(b)(1)). In conventional twin turboprop airplanes, 
this isolation is, in part, provided by the inherent separation of 
having each engine mounted on opposite sides of the airplane driving 
its own propeller. Installation of the engines on either side of the 
airplane automatically provides a degree of separation of critical 
systems, such as the electrical and fuel systems, and minimizes the 
effect of high vibration, rotor burst failures, and engine case burn-
through from the opposite engine. This separation aids in preventing 
any single failure from jeopardizing continued safe operation of the 
airplane. In contrast, the nearness of the engines to each other 
driving a combining gearbox with a single propeller in the Model LM 200 
airplane arrangement is inherently less isolated from certain types of 
failure modes. As a result, many failure modes that do not pose a 
significant hazard on conventional multi-engine airplanes could 
threaten continued safe operation of the Model LM 200 airplane unless 
specific additional precautions are taken to prevent hazardous 
secondary effects.
    The FAA has reviewed the part 23 standards and identified that 
Secs. 23.53(c), 23.67(c), 23.69, 23.75, and 23.77 are inadequate to 
address the effects of propeller control system failure modes in a 
manner consistent with how these sections address specific engine 
failure conditions. Sections 23.1191(a) and 23.1191(b) do not 
adequately define the locations of firewalls needed to isolate the 
engines and CGB of the PSU. Additionally, the FAA has identified that 
Sec. 23.1305(c) is inadequate because it does not recognize the 
uniqueness of the Model LM 200 PSU. Furthermore, the FAA has identified 
that Secs. 23.1583(b), 23.1585(c), and 23.1587(a) do not recognize a 
propeller system installation independent from either engine. Elements 
of these proposed special conditions have been developed to ensure that 
these unique aspects of the Model LM 200 airplane are addressed in a 
manner equivalent to that established by part 23 standards. The FAA's 
analysis and derivation of each of the special condition requirements 
is discussed in the Description of Proposed Requirements section below.

Description of Proposed Requirements

    The Model LM 200 will incorporate the following novel or unusual 
design features:

(a) PSU Reliability

    In order to define special conditions with the goal of establishing 
a safety level acceptable for certification as a limited commuter 
category airplane, the unique configuration of the Model LM 200 single 
propeller, twin engine design must be addressed. The Model LM 200 PSU 
design has eliminated as many single point failures as feasible for 
this type of configuration; however, certification criteria for the 
remaining single point failures unique to this configuration must be 
considered. A System Safety Analysis of the PSU is proposed that will 
identify and classify all possible failures that could be hazardous or 
catastrophic to the Model LM 200. The System Safety Analysis will 
consider such factors as non-redundancy, quality of manufacture and 
maintenance for continued airworthiness, as well as anticipated human 
errors, and it will highlight critical procedures that should be 
considered as required inspection items. Parts identified in the PSU 
System Safety Analysis whose failure results in a hazardous or 
catastrophic event will require control via a Critical Parts Plan. 
Furthermore, critical failure modes that could result in hazardous or 
catastrophic events should be addressed with appropriate design 
features to mitigate the potential results of such events.
    The critical parts plan should be modeled after plans required by 
14 CFR part 29, Sec. 29.602, and related advisory material in Advisory 
Circular 29-2C for critical rotorcraft components. In addition, best 
industry practices shall be utilized in the definition and 
implementation of these critical parts. This plan will draw the 
attention of the personnel involved in the design, manufacture, 
maintenance, and overhaul of a critical part to the special nature of 
the part. The plan should define the details of relevant special 
instructions to be included in the Instructions for Continued 
Airworthiness. The Instructions for Continued Airworthiness, required 
by Sec. 23.1529 should contain, as appropriate, life limits, mandatory 
overhaul intervals, enhanced inspection limits, periodic ultrasonic (or 
equivalent) inspections, enhanced annual inspections, and conservative 
damage limits for return to service and repair for the critical parts 
identified in accordance with these proposed special conditions.
    A means of annunciating hazardous and catastrophic failures to the 
cockpit should be provided if they are not immediately identifiable to 
the flight crew. Appropriate inspection intervals must be proposed to 
address any possible latent failures, which may go undetected.
    For those failure modes unique to the non-conventional Model LM 200 
design, which have a fail-safe designed backup, either an acceptable 
test or analysis, or both, must address worst case conditions to 
substantiate the design.

[[Page 49516]]

Methods to periodically check the backup system shall also be provided, 
as appropriate. In addition, a means of annunciating failure of the 
primary to the cockpit should be provided if it is not immediately 
identifiable to the flight crew. Appropriate inspection intervals must 
be proposed to address any possible latent failures, which may go 
undetected.

(b) Powerplant Requirements

    Although rare, high-energy rotor unbalances due to high energy 
rotating machinery failures, such as a rim separation, can occur in-
flight. They are typically followed quickly by either an in-flight 
shutdown or a pilot-commanded engine shutdown. The proposed special 
conditions address this short duration following a rotor failure by 
requiring that any high-energy vibration not affect the airworthiness 
of the operating engine. These vibrations could otherwise affect the 
operating engine in areas such as rotation (rubs), compressor surge or 
stall, damage to engine controls, accessories, mechanical, lubrication, 
fuel systems, and possible engine misalignment with respect to the 
gearbox. The magnitudes, frequency, and duration of such a vibration 
should be included in the powerplant installation manual. In addition, 
the vibration should not affect the structural integrity of the 
mounting system of either engine or the combining gearbox.
    The CGB includes all parts necessary to transmit power from the 
engines to the propeller shaft. This includes couplings, supporting 
bearings for shafts, brake assemblies, clutches, gearboxes, 
transmissions, any attached accessory pads or drives, and any cooling 
fans that are attached to, or mounted on, the gearbox. The CGB for this 
multi-engine installation must be designed with a ``continue to run'' 
philosophy. This means that it must be able to power the propeller 
after failure of one engine or failure in one side of the CGB drive 
system, including any gear, bearing, or element expected to fail. 
Common failures, such as oil pressure loss or gear tooth failure, in 
the CGB must not compromise power output from the propulsion system.
    Current engine certification regulations do not adequately address 
the requirements of a single combining gearbox; therefore, in addition 
to the engine requirements of Sec. 23.903, the CGB will be required to 
complete a 200 hour endurance test that is patterned after the rotor 
drive system requirements of Sec. 29.923. The endurance test is 
intended to exercise integration of the engines, combining gearbox and 
loading characteristics of the intended propeller. Additional testing 
patterned after Sec. 29.927 will address the torque and speed limits. 
The CGB design, should contain features that include automatic 
disengagement of any failed engine (reference Sec. 29.917(c)(3)), 
independent lubrication systems (reference Sec. 29.1027), indicators to 
alert the pilot of lubrication system failure, and the capability to 
continue safe flight to a landing for a minimum of one-hour following 
pilot notification of primary lubrication system failure.
    The requirement for continued safe flight to a landing for a 
minimum of one-hour following pilot notification of primary lubrication 
system failure stems from similarities between the Model LM 200 
propulsion system and that of a typical multi-engine rotorcraft. 
Transport category A rotorcraft must be capable of sustaining flight 
for 30-minutes after the crew is notified of a drive system lubrication 
system failure or loss of lubricant, Sec. 29.927(c). A rotorcraft may 
autorotate to a small landing area and, therefore, may find a safe 
landing area much sooner than a 19,000 pound airplane. For this reason, 
the FAA is similarly proposing that the Model LM 200 demonstrate its 
ability to sustain flight for one-hour, in accordance with AFM 
instructions for an emergency landing, after crew notification of a 
lubrication failure.
    The critical parts of the CGB must also undergo a fatigue 
evaluation patterned after the structural requirements of Sec. 29.571 
for transport rotorcraft.
    The FAA proposes the CGB should have an Initial Maintenance 
Interval established similar to the requirements for an engine in 
Sec. 33.90. The Initial Maintenance Interval will be determined 
following the completion of the 200 hour CGB endurance test and other 
proposed CGB tests.
    A rotor disc fragment should not be allowed to compromise the 
structural integrity of the powerplant or engine mounts. Loss of the 
structural integrity of the powerplant mount would be considered 
catastrophic for the Model LM 200 design. The powerplant and engine 
mount principal structural elements should be fail-safe if they could 
be severed during an uncontained engine failure. All other principal 
structural elements of the powerplant and engine mounting system should 
be either fail-safe or damage tolerant.

(c) Propeller Installation

    With a multi-engine, single propeller installation, the non-
redundancy of the propeller system components from the propeller shaft 
forward becomes quite significant. In the case of the Model LM 200, 
Ayres Corporation must design against the possibility of a propeller-
related failure that could result in catastrophic loss of the airplane. 
To accomplish this task, Ayres Corporation must substantiate the 
structural integrity of their design and must establish a critical 
parts program and a continued airworthiness maintenance and inspection 
program that ensures that the propeller is maintained in an acceptable 
manner.
    The Model LM 200 airplane's single propeller system must be 
installed and maintained in such a manner as to substantially reduce or 
eliminate the occurrence of failures that would preclude continued safe 
flight and landing. To ensure the propeller installation, production 
and maintenance programs are sufficient to achieve a high level of 
reliability, these proposed special conditions include a 2,500 cycle 
validation test based on enhanced requirements of Sec. 35.41(c). The 
2,500 cycles correspond to the FAA's estimated annual usage for a 
turboprop airplane in commercial service. An airplane cycle includes 
idle, takeoff, climb, cruise, and descent. The test must utilize 
production parts installed on the powerplant and should include a wide 
range of ambient and wind conditions, several full stops, and 
validation of scheduled and unscheduled maintenance practices. The 
purpose of this test is to evaluate the system for service wear 
conditions and start/stop cycles. It is not intended to test the 
propeller vibratory loads. This evaluation may be accomplished on the 
airplane in a combination of ground and flight cycles or on a ground 
test facility. If the testing is accomplished on a ground test 
facility, the test configuration must include the PSU and all 
sufficient airframe interfacing system hardware to simulate the actual 
airplane installation and operation.
    On a conventional multi-engine airplane, the flight crew will 
secure an engine and feather the propeller to minimize effects of 
propeller imbalance. Propeller imbalance could be caused by blade 
failures or by propeller system failures such as loss of a de-icing 
boot, malfunction of a de-icing boot in icing conditions, an oil leak 
into a blade butt, asymmetric blade pitch, or a failure in a 
counterweight attachment. The Model LM 200 airplane design does not 
provide any means to reduce the vibration produced by an unbalanced 
propeller. Therefore, these proposed special conditions require that 
the engines, CGB, powerplant and engine mounting system, primary 
airframe

[[Page 49517]]

structure, and critical systems be designed to function safely in the 
high vibration environment generated by these less severe propeller 
failures. Ayres Corporation must specify the maximum allowable 
propeller unbalance. This is the maximum unbalance that will not cause 
damage to the engines, powerplant and engine mounting system, CGB, 
primary airframe structure, or to any other critical equipment that 
would jeopardize the continued safe flight and landing of the airplane. 
The vibration level caused by this unbalance must not jeopardize the 
flight crew's ability to continue to operate the airplane in a safe 
manner. Any part (or parts) whose failure (or probable combination of 
failures) would result in a propeller unbalance greater than the 
defined maximum would also be classified as a critical part.
    It should be shown by a combination of tests and analyses that the 
airplane is capable of continued safe flight and landing with the 
maximum propeller unbalance including collateral damage caused by the 
unbalance event.
    The evaluation should show that, during continued operation for one 
hour with the declared maximum unbalance, the induced vibrations will 
not cause damage either to the primary structure of the airplane or to 
critical equipment that would jeopardize continued safe flight and 
landing. The degree of flight deck vibration should not prevent the 
flight crew from operating the airplane in a safe manner. This includes 
the ability to read and accomplish checklist procedures. This 
evaluation should consider the effects on continued safe flight and 
landing from the possible damage to primary structure, including, but 
not limited to, engine mounts, inlets, nacelles, wing, and flight 
control surfaces. Consideration should also be given to the effects of 
vibratory loads on critical equipment (including connectors) mounted on 
the engine or airframe.
    The FAA understands that in the unique design of the Model LM 200 
CGB, reverse rotation of the propeller on the ground would engage the 
sprag clutch. This, in turn, would drive both engines without 
lubrication of the engine bearings or gearbox causing possible damage 
to those elements; therefore, a means must be provided to prevent any 
adverse effects resulting from propeller ``wind-milling'' on the 
ground.
    The Hamilton Sundstrand Model 586F-11 propeller meets special 
conditions imposed during the propeller type certification program 
(Docket Nos. 94-ANE-60 and 94-ANE-61). The propeller special conditions 
addressed electronic propeller and pitch control systems, a four-pound 
bird strike, lightning strike and fatigue. If the propeller had not 
been required to meet those conditions during its type certification 
program, the FAA would have required similar measures in these Model LM 
200 special conditions since the propeller is an especially critical 
component on this airplane. To meet the airplane requirements for the 
Model LM 200, the Instructions for Continued Airworthiness may need to 
be modified.

(d) Propeller Control System

    For this propeller control system, no probable multiple failures 
were identified that create a hazardous condition, therefore, these 
special conditions were written to consider single point failures in 
the primary propeller control system only.
    These proposed special conditions require the propeller control 
system to be independent of the engines such that a failure of any 
engine or the engine's control system will not result in failure or 
inability to control the propeller.
    Ayres Corporation plans to address these special conditions by 
providing a mechanical high pitch stop, which would be set to a ``get 
home'' pitch position, thereby preventing the propeller blades from 
rotating to a feather pitch position when oil pressure is lost in the 
propeller control system. This would allow the propeller to continue to 
produce a sufficient level of thrust as a fixed pitch propeller.
    In the event the propeller undergoes an uncommanded pitch change, 
these proposed special conditions require that the Model LM 200 
airplane not be placed in an unsafe condition. They also require that 
an indication of the failure be provided to the flight crew.

(e) PSU Instrumentation

    On a conventional multi-engine airplane, the pilot has positive 
indication of an inoperative engine created by the asymmetric thrust 
condition. Because of the centerline thrust of the Model LM 200 
airplane propulsion system installation, the airplane will not yaw when 
an engine or a portion of the CGB fails. The flight crew will have to 
rely on other means to determine which engine or CGB element has failed 
in order to secure the correct engine. Therefore, these proposed 
special conditions require that a clear indication of an inoperative 
engine or a failed portion of the CGB must be provided. This is 
necessary to preclude confusion by the flight crew in reacting to the 
failure and when taking appropriate action to secure the airplane in a 
safe condition for continued flight.
    Section 23.1305 requires instruments for the fuel system, engine 
oil system, fire protection system, and propeller control system. This 
rule is intended for powerplants consisting of a single engine, 
gearbox, and propeller. To protect the portions of the PSU that are 
independent of the engines, additional instrumentation, including 
gearbox oil pressure, oil quantity, oil temperature, propeller speed, 
propeller blade angle, engine torque, and chip detection, are required.

(f) Fire Protection, Extinguishing, and Ventilation Systems

    On a conventional twin engine airplane, the engines are 
sufficiently separated to essentially eliminate the possibility of a 
fire spreading from one engine to another. In the Model LM 200, the 
engines are in close proximity, separated only by a ballistic shield 
and firewall. The fire protection system of the Model LM 200 airplane 
must include features to isolate each fire zone from any other zone and 
the airplane in order to maintain isolation of the engines and CGB 
during a fire. Therefore, these proposed special conditions mandate 
that the firewall required per Sec. 23.1191 be extended to provide 
firewall isolation between either engine and the CGB. Furthermore, 
these special conditions require that, if the potential for fire exists 
in the CGB compartment, enough fire-extinguishing agents be available 
to supply the CGB compartment and one engine compartment with the CGB 
on a dedicated system. These proposed special conditions require that 
heat radiating from a fire originating in any fire zone must not affect 
components in adjacent compartments in such a way as to endanger the 
airplane. If the potential for fire does not exist within the CGB 
compartment, this must be substantiated by analysis
    Each fire zone should be ventilated to prevent the accumulation of 
flammable vapors. In addition, it must be designed such that it will 
not allow entry of flammable fluids, vapors, or flames from other fire 
zones. It should also be designed such that it does not create an 
additional fire hazard from the discharge vapors.

(g) Airplane Performance

    Propeller control system failures may not be catastrophic in a 
conventional commuter category airplane; however, these types of 
failures should be demonstrated as not being catastrophic for the Model 
LM 200. To ensure a comparable level of safety to

[[Page 49518]]

conventional commuter category airplanes in the event of a propeller 
control system failure, these proposed special conditions require that 
the Model LM 200 propulsion system be designed such that the airplane 
meets the one-engine-inoperative performance requirements of 
Secs. 23.53, 23.67, 23.69, 23.75 and 23.77(c) with the propeller 
control system failed placing the propeller in the most critical thrust 
producing condition with both engines operating normally.

(h) Airplane Flight Manual

    In accordance with the exemption to Sec. 23.3(d), the limitations 
section of the Airplane Flight Manual will limit the airplane to a 
maximum of 9 passengers.
    Sections 23.1583, 23.1585 and 23.1587 require pertinent information 
to be included in the Airplane Flight Manual. These rules are not 
adequate to address critical propeller failures or propeller control 
system failures on the Model LM 200 airplane. As a result, these 
proposed special conditions require that the critical procedures and 
information required by Secs. 23.1583(b), 23.1583(c), 23.1585(a), 
23.1585(c) and 23.1587(d) include consideration of these critical 
propeller failures or propeller control system failures in order to 
ensure a high level of safety for this airplane.

(i) Suction Defueling

    The Model LM 200 design includes a suction defuel capability not 
envisaged when part 23 was developed. It is understood that suction 
defuel is a common feature in part 25 airplanes. The Model LM 200 
airplane will have pressure fuel and defuel capability. Pressure 
defueling essentially entails reversing the pumps on the fueling 
vehicle and ``evacuating'' fuel under vacuum from the airplane through 
the servicing port. Section 23.979 addresses pressure fueling but not 
suction defueling. Any suction defueling components, in addition to 
meeting the general requirements for part 23 fuel systems, must also 
function as intended.

(j) FADEC Installation

    Each of the engines will be controlled by a fully redundant full 
authority digital electronic control (FADEC). Each engine will utilize 
two single channel FADEC's yielding a total of four to service the PSU. 
Each FADEC is identical containing engine and propeller control 
capability; however, only two of the four units are wired to control 
the propeller. Cross-FADEC communication provides automatic enabling of 
the automatic power reserve in case of a single engine failure during 
takeoff. During normal operation, one FADEC of each engine controls 
that engine's operation while the second FADEC remains in hot standby 
mode, with the outputs deactivated, waiting to assume control. If the 
controlling unit fails, the unit in standby mode should instantly 
assume control of the engine and propeller (if applicable), without 
noticeable discontinuity.
    As the sole means of controlling the engine and the primary means 
of controlling the propeller on the Model LM 200 airplane, the FADEC 
installation must comply with the system installation requirements of 
Sec. 23.1309. While this rule was not developed to address the 
specifics of a FADEC installation, this requirement is consistent with 
the rule's intent to cover all complex electronic systems that perform 
critical functions.

Applicability

    As discussed above, these special conditions are applicable to the 
Model LM 200. Should Ayres Corporation apply at a later date for a 
change to the type certificate to include another model incorporating 
the same novel or unusual design feature, these special conditions 
would apply to that model as well under the provisions of 
Sec. 21.101(a)(1).

Conclusion

    This action affects only certain novel or unusual design features 
of the Ayres Corporation Model LM 200 airplanes. It is not a rule of 
general applicability, and it affects only the applicant who applied to 
the FAA for approval of these features on the airplane.

List of Subjects in 14 CFR Part 23

    Aircraft, Aviation safety, Signs and symbols.

Citation

    The authority citation for these special conditions is as follows:

    Authority: 49 U.S.C. 106(g), 40113 and 44701; 14 CFR 21.16 and 
21.17; and 14 CFR 11.28 and 11.29(b).

The Proposed Special Conditions

Definitions

    For purposes of this certification program and subsequent special 
conditions, the following definitions will apply:
    Powerplant--The LHTEC model CTP800-4T powerplant, consists of two 
CTS800 derivative turboshaft engines, a GKN Westland combining gearbox 
(CGB), and the engine assembly support structure. The powerplant is 
capable of providing 2,700 shp combined output power at takeoff and 
1,350 shp with one engine inoperative. The CTP800-4T powerplant will 
obtain a 14 CFR part 33 type certificate identifying the powerplant as 
a ``twin power section turboshaft assembly.''
    Engine--An LHTEC CTS800 derivative, non-regenerative, front drive, 
free turbine power section, which includes compressor, combustor, 
turbine and accessories group. Each engine of the CTP800-4T is 
separately controlled by a fully redundant full authority digital 
electronic control (FADEC). The two engines will only be certified as 
part of the CTP800-4T powerplant. The CTP800-4T type certificate data 
sheet will include ratings and limitations for each engine in addition 
to that of the powerplant.
    Engine Assembly Support Structure--The supporting structure that 
connects the two engines to the CGB. This structure will be 14 CFR part 
33 certified as part of the CTP800-4T powerplant.
    Propulsion System Unit (PSU)--The LHTEC Model CTP800-4T powerplant 
plus the airframe-mounted non-integrated lubrication system components, 
which include the CGB oil tank and CGB/engine oil cooler as well as a 
single Hamilton Sundstrand 568F-11 propeller system.
    Combining Gearbox (CGB)--All components necessary to transmit power 
from the engines to the propeller. This includes couplings, supporting 
bearings for shafts, brake assemblies, clutches, gearboxes, 
transmissions, any attached accessory pads or drives, and any cooling 
fans that are attached to, or mounted on, the gearbox. The CGB will be 
14 CFR part 33 certified as part of the CTP800-4T powerplant.
    Multi-Engine--For the Model LM 200 and its powerplant 
configuration, ``multi-engine'' refers to the twin engine capability 
and ratings of the CTP800-4T powerplant in regard to type certification 
in the commuter category and flight operations.
    One Engine Inoperative (OEI)--For the Model LM 200 airplane, ``one 
engine inoperative'' refers to a condition in which one engine of the 
CTP800-4T powerplant is not operational and the operation of the 
propeller is unchanged.
    Accordingly, the Federal Aviation Administration (FAA) proposes the 
following special conditions as part of the type certification basis 
for the Ayres Corporation Model LM 200 airplanes.
1. PSU Reliability
    (a) A PSU System Safety Analysis is required and must identify all 
hazardous or catastrophic failures associated with the unique design of 
the PSU. The analysis must consider factors

[[Page 49519]]

such as lack of redundancy, quality of manufacture and maintenance for 
continued airworthiness, including consideration of anticipated human 
errors. Critical procedures must be identified for consideration as 
required inspection items.
    (b) Critical part failures identified in the PSU System Safety 
Analysis, which result in hazardous or catastrophic events on the 
airplane, shall be controlled via a Critical Parts Plan. The Critical 
Parts Plan must be established to ensure that each critical part is 
designed and then controlled through manufacture and maintained 
throughout its service life by the following:
    (1) Enhanced procurement and manufacturing techniques,
    (2) Continued airworthiness requirements,
    (3) Conservative life limits.
    Additionally, best industry practices shall be utilized in the 
definition and implementation of these critical parts.
    (c) Critical failure modes identified in the PSU System Safety 
Analysis, which could occur due to the indirect failure of a component 
or system, should be addressed with appropriate design features to 
mitigate the potential results of such events.
    (d) An appropriate inspection interval and instructions shall be 
established for any possible latent failure of fail-safe backup 
components.
    (e) All fail-safe designs must be approved by test or analysis 
under the most adverse operational conditions and failure modes. A 
means of annunciating failure of the primary system, which could affect 
the safe operation of the airplane, must be provided to the pilot or 
maintenance crew.
2. Powerplant Requirements
    (a) Vibration.
    (1) It must be demonstrated by analysis, test, or combination 
thereof, that high-energy rotating turbomachinery failures that create 
high-energy rotor unbalance should not affect the operation of the CGB, 
the healthy engine by vibration transmitted through the CGB, the 
integrity of the airframe, powerplant, engine mounts, or the engine 
assembly support structure and attachments, or prevent continued safe 
flight and landing.
    (2) High-energy fragment and fire shielding and surrounding engine 
structure and attachments, if attached to the engine, should be 
included in the rotor dynamics analysis or any test that affects the 
rotors.
    (b) CGB Design, Endurance Testing and Additional Tests.
    (1) CGB Design. The CGB must meet the requirements as set forth in 
paragraphs (b)(2)(i)(A) through (b)(2)(iv).
    (i) The CGB must incorporate a device to automatically disengage 
any engine from the propeller shaft if that engine fails.
    (ii) The oil supply for components of the CGB that require 
continuous lubrication must be sufficiently independent of the 
lubrication systems of the engine(s) to ensure operation without damage 
to the CGB, with any engine inoperative. Each independent lubrication 
system must function properly in the flight attitudes and atmospheric 
conditions in which an airplane is expected to operate.
    (iii) Torque limiting means must be provided on all accessory 
drives that are located on the CGB in order to prevent the torque 
limits established for those drives from being exceeded.
    (2) CGB Endurance Tests. Each part tested, as prescribed in this 
section, must be in serviceable condition at the end of the tests. No 
intervening disassembly that might affect these results may be 
conducted. An endurance test report explaining the test results and 
documenting the pre- and post-test wear measurements should be 
completed.
    (i) Endurance tests; general. In addition to the 150-hour 
powerplant test requirements of Sec. 33.87, the CGB must be tested as 
prescribed in paragraphs (b)(2)(ii)(B) through (b)(2)(ii)(I), for at 
least 200 hours plus the time required to meet paragraph (b)(2)(ii)(I). 
These tests must include the engines as well as the vibration and 
loading characteristics of the propeller and allowable takeoff 
imbalance tolerance. For the 200-hour portion, these tests must be 
conducted as follows:
    (A) Twenty each, ten-hour test cycles consisting of the test times 
and procedures in paragraphs (b)(2)(ii)(B) through (b)(2)(ii)(H); and
    (B) The test torque must be determined by actual powerplant 
limitations.
    (ii) Endurance tests; takeoff torque run. The takeoff torque 
endurance test must be conducted as follows with both engines operating 
at, or CGB input shafts loaded to, the same conditions:
    (A) The takeoff torque run must consist of one hour of alternating 
runs of five minutes operating at the torque and speed corresponding to 
takeoff power, and five minutes at as low a powerplant idle speed as 
practicable. This should be done with no airframe power extractions to 
produce the highest takeoff torque and lowest idle.
    (B) Deceleration and acceleration must be performed at the maximum 
rate. (This corresponds to a one-second power setting change from idle 
to takeoff and one second from takeoff to idle setting.) This should 
also be conducted with no airframe power extractions.
    (C) The time duration of all engines at takeoff power settings must 
total one hour and does not include the time at idle and the time 
required to go from takeoff to idle and back to takeoff speed.
    (iii) Endurance tests; maximum continuous run. Three hours of 
continuous operation, at the torque corresponding to maximum continuous 
power and speed, must be conducted with maximum airframe power 
extractions.
    (iv) Endurance tests; 90 percent of maximum continuous run. One 
hour of continuous operation, at the torque corresponding to 90 percent 
of maximum continuous power at maximum continuous rotational propeller 
shaft speed with maximum airframe power extractions.
    (v) Endurance tests; 80 percent of maximum continuous run. One hour 
of continuous operation, at the torque corresponding to 80 percent of 
maximum continuous power at the minimum rotational propeller shaft 
speed intended for this power with maximum airframe power extractions.
    (vi) Endurance tests; 60 percent of maximum continuous run. Two 
hours of continuous operation, at the torque corresponding to 60 
percent of maximum continuous power at the minimum rotational propeller 
shaft speed intended for this power with maximum airframe power 
extractions.
    (vii) Endurance tests; engine malfunctioning run. It must be 
determined whether malfunctioning of components, such as the engine 
fuel or ignition systems, or unequal engine power distribution can 
cause dynamic conditions detrimental to the drive system. If so, a 
suitable number of hours of operation must be accomplished under those 
conditions, one hour of which must be included in each cycle and the 
remaining hours of which must be accomplished at the end of 20 cycles. 
This testing is to be divided between the following four conditions by 
alternating between cycles: (1) engine #1 ``ON''/engine #2 ``IDLE''; 
(2) engine #1``ON''/engine #2 ``OFF''; (3) engine #1 ``IDLE''/engine #2 
``ON''; (4) engine #1 ``OFF''/engine #2 ``ON''. If no detrimental 
condition results, an additional hour of operation in compliance with 
paragraph (B) of this section must be conducted. This will require 100 
percent transfer of the airframe air, electrical, and hydraulics to the 
operating engine

[[Page 49520]]

within approved Installation Manual limitations.
    (viii) Endurance tests; overspeed run. One hour of continuous 
operation must be conducted at the torque corresponding to maximum 
continuous power, and at 110 percent of rated maximum continuous 
rotational propeller shaft speed. This should be performed without 
airframe power extractions for highest speed. If the overspeed is 
limited to less than 110 percent of maximum continuous speed by the 
speed and torque limiting devices, the speed used must be the highest 
speed allowable, assuming that speed and torque limiting devices, if 
any, function properly.
    (ix) Endurance tests; one-engine-inoperative application. A total 
of 160 full differential power applications must be made at takeoff 
torque and RPM. If, during these tests, it is found that a critical 
dynamic condition exists, an investigative assessment to determine the 
cause shall be performed throughout the torque/speed range. In each of 
the 160 power setting cycles (160 per engine) a full differential power 
application must be performed. In each cycle, the transition from 
clutch engagement to disengagement must occur at the critical condition 
for clutch and shaft wear.
    (3) Additional CGB Tests. Following the 200-hour endurance test, 
and without any intervening major disassembly, additional dynamic, 
endurance, and operational test and vibratory investigations must be 
performed to determine that the drive mechanism is safe. The following 
additional tests and conditions apply:
    (i) If the torque output of both engines to the CGB can exceed the 
highest engine or CGB torque limit, the following tests must be 
conducted. Under conditions with both engines operating, apply 200 
cycles to the CGB for 10 seconds each of an input torque that is at 
least equal to the lesser of--
    (A) The maximum torque used in complying with paragraph 
(b)(2)(iii)(B) plus 10 percent; or
    (B) The maximum torque attainable under normal operating 
conditions, assuming that any torque limiting devices function 
properly.
    (ii) With each engine alternately inoperative, apply the maximum 
transient torque attainable under normal operating conditions, assuming 
that any torque limiting devices function properly. Each CGB input must 
be tested at this maximum torque for at least one hour.
    (iii) The CGB must be subjected to 50 overspeed runs, each 30 plus 
or minus 3 seconds in duration, at a speed of at least 110 percent of 
maximum continuous speed, or other maximum overspeed that is likely to 
occur, plus a margin of speed approved by the Administrator for that 
overspeed condition. These runs must be conducted as follows:
    (A) Overspeed runs must be alternated with stabilizing runs from 1 
to 5 minutes duration, each 60 to 80 percent of maximum continuous 
speed.
    (B) Acceleration and deceleration must be accomplished in a period 
no longer than 10 seconds, and the time for changing speeds may not be 
deducted from the specified time for the overspeed runs.
    (iv) Each part tested, as prescribed in this section, must be in 
serviceable condition at the end of the tests. No intervening 
disassembly that might affect test results may be conducted.
    (v) If drive shaft couplings are used and shaft misalignment or 
deflections are probable, loads must be determined in establishing the 
installation limits affecting misalignment. These loads must be 
combined to show adequate fatigue life.
    (vi) The CGB must be able to continue safe operation, although not 
necessarily without damage, at a torque and rotational speed prescribed 
by the applicant that is determined to be the most critical of the 
anticipated flight conditions for at least one hour after perception by 
the flight crew of the CGB lubrication system failure or loss of 
lubricant. The demonstrated torque and rotational speed must be 
included in the instruction manual for installing and operating the 
engine required in 14 CFR part 33, Sec. 33.5.
    (4) Initial Maintenance Interval. An Initial Maintenance Interval 
(reference Sec. 33.90) for the CGB shall be determined following 
completion of the testing required by sections (b)(2)(ii) through 
(b)(2)(iii).
    (5) Fatigue Evaluation. The critical parts of the CGB must be shown 
by analysis supported by test evidence and, if available, service 
experience to be of fatigue tolerant design. The fatigue tolerance 
evaluation must include the requirements of either paragraph 
(b)(2)(v)(A), (B), or (C) of this section, or a combination thereof, 
and must include a determination of the probable locations and modes of 
damage caused by fatigue, considering environmental effects, intrinsic/
discrete flaws, or accidental damage. Compliance with the flaw 
tolerance requirements of paragraph (b)(2)(v)(A) or (B) of this section 
is required unless the applicant establishes that these fatigue flaw 
tolerant methods for a particular part cannot be achieved within the 
limitations of geometry, inspectability, or good design practice. Under 
these circumstances, the safe-life evaluation of paragraph (C) of this 
section is required.
    (i) Flaw tolerant safe-life evaluation. It must be shown that the 
critical part, with flaws present, is able to withstand repeated loads 
of variable magnitude without detectable flaw growth for the following 
time intervals--
    (A) Life of the airplane; or
    (B) Within a replacement time furnished in the Instructions for 
Continued Airworthiness.
    (ii) Fail-safe (residual strength after flaw growth) evaluation. It 
must be shown that the critical part after a partial failure is able to 
withstand design limit loads without failure within an inspection per 
the Instructions for Continued Airworthiness. Limit loads are defined 
in Sec. 23.301(a).
    (A) The residual strength evaluation must show that the critical 
part after flaw growth is able to withstand design limit loads without 
failure within its operational life.
    (B) Inspection intervals and methods must be established as 
necessary to ensure that failures are detected prior to residual 
strength conditions being reached.
    (C) If significant changes in structural stiffness or geometry, or 
both, follow from a structural failure or partial failure, the effect 
on flaw tolerance must be further investigated.
    (iii) Safe-life evaluation. It must be shown that the critical part 
is able to withstand repeated loads of variable magnitude without 
detectable cracks for the following time intervals--
    (A) Life of the airplane; or
    (B) Within a replacement time furnished in the Instructions for 
Continued Airworthiness.
    (C) Powerplant and Engine Mounts.
    (1) All principal structural elements of the powerplant and engine 
mount structure that could fail as a result of an uncontained engine 
failure or resulting fire must be fail-safe as defined in 
Sec. 23.571(b). All other principal structural elements of the 
powerplant and engine mount system must either be fail-safe or meet the 
damage tolerance criteria of Sec. 23.574(a).
    (i) For fail-safe design:
    (A) The fail-safe structure must be able to withstand the limit 
loads, considered as ultimate, given in Secs. 23.361 and 23.363.
    (B) If the occurrence of load-inducing propeller control systems 
malfunctions is less frequent than 1  x  10\-5\ occurrences per flight 
hour, and if it can be demonstrated that failure or partial

[[Page 49521]]

failure of a structural element would be obvious, the engine torque 
loads of Sec. 23.361(a)(3) do not need to be considered in the fail-
safe design.
    (ii) If damage tolerance evaluation is used,
    (A) The residual strength evaluation must consider the limit loads, 
considered as ultimate, given in Secs. 23.361 and 23.363.
    (B) If the occurrence of load-inducing propeller control system 
malfunctions is less frequent than 1  x  10\-5\ occurrences per flight 
hour, the engine torque loads of Sec. 23.362(a)(3) do not need to be 
considered in the residual strength evaluation.
3. Propeller Installation
    (a) The applicant must complete a 2,500 airplane cycle evaluation 
of the propeller installation. A cycle must include the power levels 
associated with ground idle, takeoff, climb cruise, and descent. This 
evaluation may be accomplished on the airplane in a combination of 
ground and flight cycles or on a ground test facility. If the testing 
is accomplished on a ground test facility, the test configuration must 
include sufficient interfacing system hardware to simulate the actual 
airplane installation, including the engines, CGB and mount system. 
Each part tested, as prescribed in this section, must be in serviceable 
condition at the end of the tests. No intervening disassembly, other 
than normal maintenance (as defined for the installation), that might 
affect these results may be conducted. A test report explaining the 
test results and documenting the pre-and post-test condition should be 
completed.
    (b) Propeller Unbalance. It must be shown by a combination of 
testing and analysis that any single failure or probable combination of 
failures, not deemed a critical part under paragraph (a)(4), that could 
cause an unbalanced propeller condition will not cause damage to the 
engines, CGB, powerplant mount system, primary airframe structure, or 
to critical equipment that would jeopardize the continued safe flight 
and landing of the airplane. Furthermore, the degree of flight deck 
vibration must not jeopardize the crew's ability to continue to operate 
the airplane in a safe manner. The magnitude and frequency of the 
vibration should be included in the installation manual. Any part (or 
parts) whose failure (or combination of failures) would result in a 
propeller unbalance greater than the defined maximum should also be 
classified as critical.
    (c) A means must be provided to prevent any adverse effect 
resulting from rotation of the propeller, in either direction, on the 
ground.
4. Propeller Control System
    (a) The propeller control must be independent of the engines, such 
that a failure in either engine or any engine control system will not 
result in failure to control the propeller.
    (b) The propeller control system must be designed to minimize the 
occurrence of any single failure that would prevent the propulsion 
system from producing thrust at a level required to meet 
Secs. 23.53(c), 23.67(c), 23.69, 23.75, and 23.77(c).
    (c) An uncommanded propeller pitch change must not result in an 
unsafe condition and an indication of the failure must be annunciated 
to the flight crew.
5. PSU Instrumentation
    (a) Engine Failure Indication. A means must be provided to indicate 
when an engine is no longer able to provide torque, or to provide 
stable torque, to the propeller. This means may consist of 
instrumentation required by other sections of part 23 or these special 
conditions if it is determined that those instruments will readily 
alert the flight crew when a engine is no longer able to provide 
torque, or to provide stable torque, to the propeller. This indicator 
must preclude confusion by the flight crew in reacting to the failure 
and when taking appropriate action to secure the airplane in a safe 
condition for continued flight.
    (b) Engine/Propeller Vibration Exceedance Indication. A means must 
be provided to indicate when the PSU vibration levels exceed the 
maximum vibration level defined for continuous operation. Procedures to 
respond to this exceedance should be included in the AFM.
    (c) The engine instrumentation requirements of Sec. 23.1305 (a), 
(c) and (e) shall apply to each engine as defined in these special 
conditions.
    (d) In addition to the requirements of Sec. 23.1305, the following 
instruments must be provided:
    (1) An oil pressure warning means and indicator for the pressure-
lubricated CGB to indicate when the oil pressure falls below a safe 
value.
    (2) A low oil quantity indicator for the CGB, if lubricant is self-
contained;
    (3) An oil temperature warning device to indicate unsafe CGB 
temperatures;
    (4) A tachometer for the propeller;
    (5) A propeller pitch control failure indication;
    (6) A torquemeter for each engine if the sum of the maximum torque 
that each engine is capable of producing exceeds the maximum torque for 
which the CGB has been certified under 14 CFR part 33; and
    (7) A chip detecting and indicating system for the CGB.
6. Fire Protection, Extinguishing and Ventilation Systems
    (a) Each engine must be isolated from the other engine and CGB by 
firewalls, shrouds, or equivalent means. Each firewall or shroud, 
including applicable portions of the engine couplings, must be 
constructed such that no hazardous quantity of liquid, gas, or flame 
can pass between the isolated fire zone of each engine or the CGB 
compartment.
    (b) In addition to the engine fire zones, if the potential for fire 
exists in the CGB compartment, then the CGB must be in a separate fire 
zone and must comply with all fire protection requirements of 14 CFR 
part 23. Enough fire-extinguishing agent will be required for the CGB 
compartment and at least one engine compartment. A dedicated fire 
extinguishing system will be required for the CGB compartment. If the 
potential for fire does not exist within the CGB compartment, this must 
be substantiated by analysis.
    (c) Firewall temperatures under all normal or failure conditions 
must not result in auto-ignition of flammable fluids and vapors present 
in the other engine compartment and the CGB compartment.
    (d) The CGB compartment ventilation system must be designed such 
that:
    (1) It is ventilated to prevent the accumulation of flammable 
vapors.
    (2) No ventilation opening may be where it would allow the entry of 
flammable fluids, vapors, or flame from other zones.
    (3) Each ventilation means must be arranged so that no discharged 
vapors will cause an additional fire hazard.
    (4) Unless the extinguishing agent capacity and rate of discharge 
are based on maximum airflow through the compartment, there must be a 
means to allow the crew to shut off sources of forced ventilation.
7. Cargo or Baggage Compartment Requirements
    (a) Flight tests must demonstrate means to exclude hazardous 
quantities of smoke, flames, or extinguishing agent from any 
compartment occupied by the crew or passengers.
    (b) Cargo compartments shall have either fire or smoke detection 
provisions, or both, unless the compartment location is such that a 
fire can be easily detected by the pilots seated at their duty station. 
The cargo and baggage fire protection must be in

[[Page 49522]]

accordance with Sec. 23.855 as well as the following:
    (1) The detection system must provide a visual indication to the 
flight crew within one minute after the start of a fire.
    (2) 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.
    (3) There must be means to allow the crew to check the functioning 
of each fire detector circuit while in flight.
    (4) The detection system effectiveness must be shown for all 
approved operating configurations and conditions.
    (c) The flight crew must have means to shut off the ventilating 
airflow to, or within, the compartment from the pilot's station on the 
all-cargo configuration.
    (d) Passenger and combi configurations, where the cargo compartment 
is not accessible to the flight crew, must have an approved built-in 
fire extinguishing system. The built-in fire extinguishing system shall 
be controllable from the pilots' station. There must be means to 
control ventilation and drafts within the inaccessible cargo 
compartment so that the extinguishing agent can control any fire that 
may start within the compartment. The built-in fire extinguisher must 
be installed so that no extinguishing agent likely to enter personnel 
compartments will be hazardous to the occupants. The discharge of the 
extinguisher must not cause structural damage. The capacity of the 
extinguishing system must be adequate for any fire likely to occur in 
the compartment where used. Consideration must be given to the volume 
of the compartment and the ventilation rate.
    (e) In addition to the hand fire extinguishers required by 
Sec. 23.851, a hand fire extinguisher must be readily accessible for 
use in each cargo or baggage compartment that is accessible to 
crewmembers in flight. Hazardous quantities of smoke, flames, or 
extinguishing agent must not enter any compartment occupied by the crew 
or passengers when the access to that compartment is used.
    (f) Protective breathing equipment must be installed for 
crewmembers in each crewmember compartment. Protective breathing 
equipment must:
    (1) Be designed to protect the flight crew from smoke, carbon 
dioxide, and other harmful gases at the pilot's station and while 
combating fires in cargo compartments.
    (2) Have masks that cover the eyes, nose, and mouth; or masks that 
cover the nose and mouth plus accessory equipment to cover the eyes.
    (3) Allow the flight crew to use the radio equipment and to 
communicate with each other while at their assigned stations.
    (4) Not cause any appreciable adverse effect on vision and must 
allow corrective glasses to be worn.
    (5) 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. 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 protective breathing 
system is used (including a mask with a standard rebreather bag) 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. BTPD refers to body temperature 
conditions (that is, 37 deg. C., at ambient pressure, dry).
    (6) Be free from hazards in itself, in its method of operation, and 
in its effect upon other components.
    (7) Have a means to allow the crew to readily determine, during 
flight, the quantity of oxygen available in each source of supply.
8. Airplane Performance
    (a) In addition to the takeoff performance requirements of 
Sec. 23.53(c), the same requirements must be met with both engines 
operating normally and the propeller primary control system failed in 
the most critical thrust producing condition at VEF and 
above, considering all single point failures.
    (b) In addition to the one engine inoperative climb requirements of 
Sec. 23.67(c), the same requirements must be met with both engines 
operating normally and the propeller primary control system failed in 
the most critical thrust producing condition, considering all single 
point failures.
    (c) In addition to the requirements of Sec. 23.69, 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 both engines operating normally and 
the propeller primary control system failed in the most critical thrust 
producing condition, considering all single point failures.
    (d) In addition to Sec. 23.75, the horizontal distance necessary to 
land and come to a complete stop from a point 50 feet above the landing 
surface must be determined as required in Sec. 23.75 with both engines 
operating normally and the propeller primary control system failed in 
the most critical thrust producing conditions, considering all single 
point failures.
    (e) The balked landing requirements of Sec. 23.77(c) must be 
performed with the propeller primary control system failed in the most 
critical thrust producing condition, considering all single point 
failures.
9. Airplane Flight Manual
    (a) In addition to the requirements of Secs. 23.1583(b) and 
23.1585(a), a pre-flight visual inspection of the propeller components 
must be included in the Airplane Flight Manual.
    (b) In addition to the requirements of Sec. 23.1585(c), procedures 
for maintaining or recovering control of the airplane in all conditions 
identified in section (g) of these special conditions must be included 
in the Airplane Flight Manual.
    (c) The information required by Sec. 23.1583(c)(4) and 
Sec. 23.1587(d) must be furnished with the propeller control system 
failed or with one engine inoperative, whichever is more critical.
10. Suction Defueling
    (a) 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 maximum permissible defueling pressure (positive or 
negative) at the airplane fueling connection.
11. FADEC Installation
    (a) The installation of the electronic engine/propeller control 
(FADEC control system) must comply with the requirements of 
Sec. 23.1309(a) through (e).

    Issued in Kansas City, Missouri on July 19, 2000.
Michael Gallagher,
Manager, Small Airplane Directorate, Aircraft Certification Service.
[FR Doc. 00-20584 Filed 8-11-00; 8:45 am]
BILLING CODE 4910-13-U