[Federal Register Volume 62, Number 211 (Friday, October 31, 1997)]
[Rules and Regulations]
[Pages 58875-58890]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 97-28937]



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Rules and Regulations
                                                Federal Register
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Federal Register / Vol. 62, No. 211 / Friday, October 31, 1997 / 
Rules and Regulations

[[Page 58875]]


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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 23

[Docket No. 135CE, Special Conditions 23-ACE-87]


Special Conditions; Sino Swearingen Model SJ30-2 Airplane

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Final special conditions.

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SUMMARY: These special conditions are being issued to become part of 
the type certification basis for the Sino Swearingen Aircraft Company 
Model SJ30-2 airplane. This new airplane will have novel and unusual 
design features not addressed in the airworthiness standards for 
normal, utility, acrobatic, and commuter category airplanes. These 
design features include a high operating altitude (49,000 feet), swept 
wings and stabilizer, performance characteristics, large fuel capacity, 
and protection for the electronic engine control and flight and 
navigation systems from high intensity radiated fields, for which the 
applicable regulations do not contain adequate or appropriate 
airworthiness standards. These special conditions contain the 
additional airworthiness standards that the Administrator considers 
necessary to establish a level of safety equivalent to that existing in 
the current business jet fleet and expected by the user of this class 
of aircraft.

EFFECTIVE DATE: December 1, 1997.

FOR FURTHER INFORMATION CONTACT: Lowell Foster, Aerospace Engineer, 
Standards Office (ACE-110), Small Airplane Directorate, Aircraft 
Certification Service, Federal Aviation Administration, Room 1544, 601 
East 12th Street, Kansas City, Missouri 64106; telephone (816) 426-
5688.

SUPPLEMENTARY INFORMATION:

Background

    On October 9, 1995, Sino Swearingen Aircraft Company, 1770 Sky 
Place Boulevard, San Antonio, Texas 78216, made application for normal 
category type certification of its Model SJ30-2 airplane, a six-to-
eight place, all metal, low-wing, T-tail, twin turbofan engine powered 
airplane with fully enclosed retractable landing gear. The SJ30-2 will 
have a V<INF>MO</INF>/M<INF>MO</INF> of 320 kts/M=.83, and will have 
engines mounted aft on the fuselage.

Type Certification Basis

    Type certification basis of the Model SJ30-2 airplane is: 14 CFR 
Part 23, effective February 1, 1965, through Amendment 23-52, effective 
July 25, 1996; 14 CFR Part 36, effective December 1, 1969, through the 
amendment effective on the date of type certification; 14 CFR Part 34; 
exemptions, if any; and the special conditions adopted by this 
rulemaking action.

Discussion

    Special conditions may be issued and amended, as necessary, as part 
of the type certification basis if the Administrator finds that the 
airworthiness standards designated in accordance with 14 CFR Part 21, 
Sec. 21.17(a)(1), do not contain adequate or appropriate safety 
standards because of novel or unusual design features of an airplane. 
Special conditions, as appropriate, are issued in accordance with 14 
CFR Part 11, Sec. 11.49, after public notice, as required by 
Secs. 11.28 and 11.29(b), effective October 14, 1980, and become part 
of the type certification basis as provided by part 21, 
Sec. 21.17(a)(2).

Protection of Systems From High Intensity Radiated Fields (HIRF)

    Recent advances in technology have led to the application in 
aircraft designs of advanced electrical and electronic systems that 
perform functions required for continued safe flight and landing. Due 
to the use of sensitive solid state advanced components in analog and 
digital electronics circuits, these advanced systems are readily 
responsive to the transient effects of induced electrical current and 
voltage caused by the HIRF. The HIRF can degrade electronic systems 
performance by damaging components or upsetting system functions.
    Furthermore, the HIRF environment has undergone a transformation 
that was not foreseen when the current requirements were developed. 
Higher energy levels are radiated from transmitters that are used for 
radar, radio, and television. Also, the number of transmitters has 
increased significantly. There is also uncertainty concerning the 
effectiveness of airframe shielding for HIRF. Furthermore, coupling to 
cockpit-installed equipment through the cockpit window apertures is 
undefined.
    The combined effect of the technological advances in airplane 
design and the changing environment has resulted in an increased level 
of vulnerability of electrical and electronic systems required for the 
continued safe flight and landing of the airplane. Effective measures 
against the effects of exposure to HIRF must be provided by the design 
and installation of these systems. The accepted maximum energy levels 
in which civilian airplane system installations must be capable of 
operating safely are based on surveys and analysis of existing radio 
frequency emitters. These special conditions require that the airplane 
be evaluated under these energy levels for the protection of the 
electronic system and its associated wiring harness. These external 
threat levels, which are lower than previous required values, are 
believed to represent the worst case to which an airplane would be 
exposed in the operating environment.
    These special conditions require qualification of systems that 
perform critical functions, as installed in aircraft, to the defined 
HIRF environment in paragraph 1 or, as an option to a fixed value using 
laboratory tests, in paragraph 2, as follows:
    (1) The applicant may demonstrate that the operation and 
operational capability of the installed electrical and electronic 
systems that perform critical functions are not adversely affected when 
the aircraft is exposed to the HIRF environment defined below:

                       Field Strength Volts/Meter                       
------------------------------------------------------------------------
                       Frequency                          Peak   Average
------------------------------------------------------------------------
10-100 KHz............................................       50       50
100-500...............................................       60       60

[[Page 58876]]

                                                                        
500-2000..............................................       70       70
2-30 MHz..............................................      200      200
30-70.................................................       30       30
70-100................................................       30       30
100-200...............................................      150       30
200-400...............................................       70       70
400-700...............................................      700       80
700-1000..............................................     1700      240
1-2 GHz...............................................     5000      360
2-4...................................................     4500      360
4-6...................................................     7200      300
6-8...................................................     2000      330
8-12..................................................     3500      270
12-18.................................................     3500      330
18-40.................................................      780       20
------------------------------------------------------------------------

or,
    (2) The applicant may demonstrate by a system test and analysis 
that the electrical and electronic systems that perform critical 
functions can withstand a minimum threat of 100 volts per meter, peak 
electrical field strength, from 10 KHz to 18 GHz. When using this test 
to show compliance with the HIRF requirements, no credit is given for 
signal attenuation due to installation.
    A preliminary hazard analysis must be performed by the applicant, 
for approval by the FAA, to identify electrical and/or electronic 
systems that perform critical functions. The term ``critical'' means 
those functions whose failure would contribute to, or cause, a failure 
condition that would prevent the continued safe flight and landing of 
the airplane. The systems identified by the hazard analysis that 
perform critical functions are candidates for the application of HIRF 
requirements. A system may perform both critical and non-critical 
functions. Primary electronic flight display systems, and their 
associated components, perform critical functions such as attitude, 
altitude, and airspeed indication. The HIRF requirements apply only to 
critical functions.
    Compliance with HIRF requirements may be demonstrated by tests, 
analysis, models, similarity with existing systems, or any combination 
of these. Service experience alone is not acceptable since normal 
flight operations may not include an exposure to the HIRF environment. 
Reliance on a system with similar design features for redundancy as a 
means of protection against the effects of external HIRF is generally 
insufficient since all elements of a redundant system are likely to be 
exposed to the fields concurrently.

Performance

    The Sino Swearingen Model SJ30-2 has a main wing with 30 degrees of 
leading-edge sweepback that employs leading-edge slats and Fowler-
flaps. The airplane has a T-tail with trimmable horizontal stabilizer 
and 30 degrees of leading-edge sweepback. There are two medium bypass 
ratio turbofan engines mounted on the aft fuselage.
    Previous certification and operational experience with airplanes of 
like design in the transport category reveal certain unique 
characteristics compared to conventional aircraft certificated under 
part 23. These characteristics have caused safety problems in the past 
when pilots attempted takeoffs and landings, particularly with a large 
variation in temperature and altitude, using procedures and instincts 
developed with conventional airplanes.
    One of the major distinguishing features of a swept-wing design not 
considered in current part 23 is a characteristically flatter lift 
curve without a ``stall'' break near the maximum coefficient of lift, 
as in a conventional wing. The ``stall'' separation point may occur at 
a much higher angle of attack than the point of maximum lift and the 
angle of attack for maximum lift can be only recognized by precise test 
measurements or specific detection systems. This phenomenon is not 
apparent to a pilot accustomed to operating a conventional airplane 
where increasing angle of attack produces increased lift to the point 
where the wing stalls. In a swept-wing design, if the pilot does not 
operate in accordance with established standards developed through a 
dedicated test program, increasing angle of attack may produce very 
little lift yet increase drag markedly to the point where flight is 
impossible. These adverse conditions may be further compounded by the 
characteristics of turbofan engines, including specified N<INF>1</INF>/
N<INF>2</INF> rotational speeds, temperature, and pressure limits that 
make its variation in thrust output with changes in temperature and 
altitude more complex and difficult to predict. In recognition of these 
characteristics, Special Civil Air Regulations No. SR-422, and follow-
on regulations, established weight-altitude-temperature (WAT) 
limitations and procedures for scheduling takeoff and landing for 
turbine powered transport category airplanes, so the pilot could 
achieve reliable and repeatable results under all expected conditions 
of operation. This entails specific tests such as minimum unstick 
speed, V<INF>MU</INF>, to ensure that rotation and fly-out speeds are 
correct and that the airplane speed schedule will not allow the 
airplane to lift off in ground effect and then be unable to accelerate 
and continue to climb out. In conjunction with the development of 
takeoff and landing procedures, it was also necessary to establish 
required climb gradients and data for flight path determination under 
all approved weights, altitudes, and temperatures. This enables the 
pilot to determine, before takeoff, that a safe takeoff, departure, and 
landing at destination can be achieved.

Takeoff

    Based upon the knowledge and experience gained with similar high 
speed, high efficiency, turbojet airplanes with complex high lift 
devices for takeoff and landing, special conditions require performance 
standards for takeoff, takeoff speeds, accelerate-stop distance, 
takeoff path, takeoff distance, takeoff run, and takeoff flight path.
    Additionally, procedures for takeoff, accelerate-stop distance, and 
landing are proposed as those established for operation in service and 
must be executable by pilots of average skill and include reasonably 
expected time delays.

Climb

    To maintain a level of safety that is equivalent to the current 
business jet fleet for takeoff, takeoff speeds, takeoff path, takeoff 
distance, and takeoff run, it is appropriate to require specific climb 
gradients, airplane configurations, and consideration of atmospheric 
conditions that will be encountered. These special conditions include 
climb with one engine inoperative, balked landing climb, and general 
climb conditions.

Landing

    Landing distance determined for the same parameters is consistent 
with takeoff information for the range of weights, altitudes, and 
temperatures approved for operation. Further, it is necessary to 
consider time delays to provide for in-service variation in the 
activation of deceleration devices, such as spoilers and brakes.

Trim

    Special conditions are issued to maintain a level of safety that is 
consistent with the use of V<INF>MO</INF>/M<INF>MO</INF> and the 
requirements established for previous part 23 jet airplanes. Current 
standards in part 23 did not envision this type of airplane and the 
associated trim considerations.

[[Page 58877]]

Demonstration of Static Longitudinal Stability

    To maintain a level of safety consistent with existing business jet 
airplanes, it is appropriate to define applicable requirements for 
static longitudinal stability. Current standards in part 23 did not 
envision this type of airplane and the associated stability 
considerations. Special conditions will establish static longitudinal 
stability requirements that include a stick force versus speed 
specification and stability requirements applicable to high speed jet 
airplanes.
    Consistent with the concept of V<INF>MO</INF>/M<INF>MO</INF> being 
a maximum operational speed limit, rather than a limiting speed for the 
demonstration of satisfactory flight characteristics, it is appropriate 
to extend the speed for demonstration of longitudinal stability 
characteristics from the V<INF>MO</INF>/M<INF>MO</INF> of 14 CFR Part 
23 to the maximum speed for stability characteristics, V<INF>FC</INF>/
M<INF>FC</INF>, for this airplane.

Static Directional and Lateral Stability

    Consistent with the concept of V<INF>MO</INF>/M<INF>MO</INF> being 
a maximum operational speed limit, rather than a limiting speed for the 
demonstration of satisfactory flight characteristics, it is appropriate 
to extend the speed for demonstration of lateral/directional stability 
characteristics from the V<INF>MO</INF>/M<INF>MO</INF> of part 23 to 
the maximum speed for stability characteristics, V<INF>FC</INF>/
M<INF>FC</INF> for this airplane.
    Current transport category regulations have eliminated the 
independent lateral stability demonstration requirement (picking up the 
low wing with rudder application). This requirement was originally 
intended to provide adequate controllability in the event of lateral 
control system failure. Because the SJ30-2 flight control system 
reliability requirement is not to current transport category levels, it 
is appropriate to retain the prior transport category requirements to 
retain the independent dihedral effect and skid recovery demonstration 
requirements.

Stall Characteristics

    The stall characteristics requirements are relaxed from part 23 to 
be equivalent to that acceptable in current business jets. These 
special conditions reflect a higher expected pilot proficiency level, 
the remote chance that a stall will be encountered in normal operation, 
and are relaxed as compensation for meeting higher performance 
requirements in these special conditions.

Vibration and Buffeting

    The Sino Swearingen Model SJ30-2 will be operated at high altitudes 
where stall-Mach buffet encounters (small speed margin between stall 
and transonic flow buffet) are likely to occur, which is not presently 
addressed in part 23. The special condition will require buffet onset 
tests and the inclusion of information in the Airplane Flight Manual 
(AFM) to provide guidance to the flightcrew. This information will 
enable the flightcrew to plan flight operations that will maximize the 
maneuvering capability during high altitude cruise flight and preclude 
intentional operations exceeding the boundary of perceptible buffet. 
Buffeting is considered to be a warning to the pilot that the airplane 
is approaching an undesirable and eventually dangerous flight regime, 
that is, stall buffeting, high speed buffeting or maneuvering (load 
factor) buffeting. In straight flight, therefore, such buffet warning 
should not occur at any normal operating speed up to the maximum 
operating limit speed, V<INF>MO</INF>/M<INF>MO</INF>.

High Speed Characteristics and Maximum Operating Limit Speed

    The Sino Swearingen Model SJ30-2 will be operated at high altitude 
and high speeds. The proposed operating envelope includes areas in 
which Mach effects, which have not been considered in part 23, may be 
significant. The anticipated low drag of the airplane and the proposed 
operating envelope are representative of the conditions not envisioned 
by the existing part 23 regulations. These conditions may degrade the 
ability of the flightcrew to promptly recover from inadvertent 
excursions beyond maximum operating speeds. The ability to pull a 
positive load factor is needed to ensure, during recovery from upset, 
that the airplane speed does not continue to increase to a value where 
recovery may not be achievable by the average pilot or flightcrew.
    Additionally, to allow the aircraft designer to conservatively 
design to higher speeds than may be operationally required for the 
airplane, the concept of V<INF>DF</INF>/M<INF>DF</INF>, the highest 
demonstrated flight speed for the type design, is appropriate for this 
airplane. This permits V<INF>D</INF>/M<INF>D</INF>, the design dive 
speed, to be higher than the speed actually required to be demonstrated 
in flight. Accordingly, the special conditions allow determination of a 
maximum demonstrated flight speed and to relate the determination of 
V<INF>MO</INF>/M<INF>MO</INF> to the speed V<INF>DF</INF>/
M<INF>DF</INF>.

Flight Flutter Tests

    Flight flutter test special conditions are proposed to 
V<INF>DF</INF>/M<INF>DF</INF> rather than to V<INF>D</INF>, in keeping 
with the V<INF>DF</INF>/M<INF>DF</INF> concept.

Out-of-Trim Characteristics

    High speed airplanes have experienced a number of upset incidents 
involving out-of-trim conditions. This is particularly true for swept-
wing airplanes and airplanes with a trimmable stabilizer. Service 
experience has shown that out-of-trim conditions can occur in flight 
for various reasons and that the control and maneuvering 
characteristics of the airplane may be critical in recovering from 
upsets. The existing part 23 regulations do not address high speed out-
of-trim conditions. These special conditions test the out-of-trim 
flight characteristics by requiring the longitudinal trim control be 
displaced from the trimmed position by the amount resulting from the 
three-second movement of the trim system at this normal rate with no 
aerodynamic load, or the maximum mis-trim that the autopilot can 
sustain in level flight in the high speed cruise condition, whichever 
is greater. Special conditions require the maneuvering characteristics, 
including stick force per g, be explored throughout a specified 
maneuver load factor speed envelope. The dive recovery characteristics 
of the aircraft in the out-of-trim condition specified would be 
investigated to determine that safe recovery can be made from the 
demonstrated flight dive speed V<INF>DF</INF>/M<INF>DF</INF>.

Pressure Vessel Integrity

    Special conditions will be used to ensure pressure vessel integrity 
for operation at altitudes above 41,000 feet. The FAA uses 41,000 feet 
as the altitude where additional requirements for high altitude 
operations are necessary. Crack growth data are used to prescribe an 
inspection program that should detect cracks before an opening in the 
pressure vessel would allow rapid depressurization.

Fuel System Protection During Collapse of Landing Gear

    The SJ30-2 maximum fuel weight is 39 percent of the maximum weight. 
This percentage is typical of the turbofan powered business jet class 
of airplanes. Part 23 did not envision that the applicable airplane 
designs would have such a large fraction of maximum weight as fuel. 
Part 23 does not contain fuel system protection requirements during 
landing gear collapse, except for Sec. 23.721, which pertains to 
commuter

[[Page 58878]]

category airplanes that have a passenger seating configuration of 10 
seats or more. In the SJ30-2 design, there is a large fuselage fuel 
tank and the placement of the engines on the aft fuselage requires that 
the fuel lines be routed through the fuselage, making the fuel lines 
more vulnerable to damage, or rupture, if the landing gear collapses. 
The special condition is based on 14 CFR Part 25, Sec. 25.721(a)(1), 
which is applicable to airplanes having a passenger seating 
configuration of nine seats or fewer.

Oxygen System Equipment and Supply

    Continuous flow passenger oxygen equipment is certified for use up 
to 40,000 feet; however, for rapid decompressions above 34,000 feet, 
reverse diffusion leads to low oxygen partial pressures in the lungs to 
the extent that a small percentage of passengers may lose useful 
consciousness at 35,000 feet even with the use of the continuous flow 
system. To prevent permanent physiological damage, the cabin altitude 
must not exceed 25,000 feet for more than 2 minutes. The maximum peak 
cabin altitude of 40,000 feet is consistent with the standards 
established for previous certification programs. In addition, at high 
altitudes the other aspects of decompression sickness have a 
significant detrimental effect on pilot performance (for example, a 
pilot can be incapacitated by internal expanding gases).
    Decompression above the 37,000 foot limit depicted in Figure 4 
approaches the physiological limits of the average person; therefore, 
every effort must be made to provide the pilots with adequate oxygen 
equipment to withstand these severe decompressions. Reducing the time 
interval between pressurization failure and the time the pilots receive 
oxygen will provide a safety margin against being incapacitated and can 
be accomplished by the use of mask-mounted regulators. The proposed 
special condition, therefore, would require pressure demand masks with 
mask-mounted regulators for the flightcrew. This combination of 
equipment will provide the best practical protection for the failures 
covered by this special condition and for improbable failures not 
covered by the special conditions, provided the cabin altitude is 
limited.

Airspeed Indicating System

    To maintain a level of safety consistent with that existing in the 
current business jet fleet, and to be consistent with the establishment 
of speed schedule performance requirements, it is appropriate to 
establish applicable requirements for determining and providing 
airspeed indicating system calibration information. Additionally, it is 
appropriate to establish special conditions requiring protection of the 
pitot tube from malfunctions associated with icing conditions. Special 
conditions will establish airspeed indicating system calibration and 
pitot tube ice protection requirements applicable to transport category 
jet airplanes.

Static Pressure System

    Special conditions are appropriate to establish applicable 
requirements for providing static pressure system calibration 
information in the AFM. Since aircraft of this type are frequently 
equipped with devices to correct the altimeter indication, it is also 
appropriate to establish requirements to ensure the continued 
availability of altitude information where such a device malfunctions. 
Current standards in part 23 did not envision this type of airplane and 
the associated static pressure requirements.

Minimum Flightcrew

    The Sino Swearingen Model SJ30-2 operates at high altitudes and 
speeds not envisioned in part 23 and must be flown in a precise speed 
schedule to achieve flight manual takeoff and landing distances. 
Therefore, it is appropriate to specify workload considerations. 
Special conditions will specify the items to be considered in workload 
determination.

Airplane Flight Manual (AFM) Information

    To be consistent with the performance special conditions, it is 
also necessary to require that the maximum takeoff and landing weights, 
takeoff distances, and associated atmospheric conditions be made 
available to the pilot in the AFM and that the airplane be operated 
within its performance capabilities. Special conditions will add 
maximum takeoff weights, maximum landing weights, and minimum takeoff 
distances as limitations in the AFM. Additionally, special conditions 
are included to add takeoff flight path and procedures necessary to 
achieve the performance in the limitations section as information in 
the AFM.

Discussion of Comments

    Notice of Proposed Special Conditions, Notice No. 23-ACE-87, Docket 
No. 135CE, was published in the Federal Register on February 21, 1997, 
and the comment period closed March 24, 1997. Following is a summary of 
the comments received and a response to each comment.
    Only one commenter responded to the notice of proposed special 
conditions and that was the Sino Swearingen Aircraft Company. They 
offered 15 comments, of which 7 were either editorial in nature or the 
incorrect special condition numbers were referenced. These errors were 
corrected. The remainder of comments are addressed individually.
    1. Comment: The certification basis should be changed to part 23 
through Amendment 23-52.
    FAA Response: The FAA agrees and the type certification basis for 
the special condition has been changed accordingly.
    2. Comment: In the discussion material section, remove the words 
``double slotted'' from the first sentence of the ``Performance'' 
discussion on page 7951, third column, first paragraph.
    FAA Response: The FAA agrees and has removed the words.
    3. Comment: Add the following statement to the ``Discussion'' 
material:

Demonstration of Static Longitudinal Stability

    To maintain a level of safety consistent with that applied to 
previous part 23 jet airplanes, it is appropriate to define applicable 
requirements for static longitudinal stability. Current standards in 
part 23 did not envision this type of airplane with the associated 
stability considerations. Special conditions are proposed to establish 
static longitudinal stability requirements that include a stick force 
versus speed specification and stability requirements applicable to 
high speed jet airplanes.
    FAA Response: The FAA concurs and has incorporated this comment 
into the section.
    4. Comment: Special Condition No. 1 lacks specificity. The 
discussion material includes the two options that we may use to show 
compliance, but the proposed special condition is silent. Suggest that 
these options be included in the body of the special condition and not 
left in the discussion material.
    FAA Response: This is the format used for HIRF special conditions. 
The FAA's goal is rules that contain minimum standards and not means of 
showing compliance. While this is hard to accomplish in certain 
instances, it is not the FAA's intention to dictate designs to 
manufacturers, but to offer compliance options through advisory 
circular. In this case, the HIRF minimum standards are the special 
conditions, which constitute a rule, and

[[Page 58879]]

one acceptable means of showing compliance is discussed in the 
preamble.
    5. Comment: Special Condition No. 24, Out-of-Trim Characteristics. 
The opening statement should be changed to ``the following applies'' 
instead of ``the Sino Swearingen model SJ30-2 must comply with the 
following.''
    FAA: The FAA agrees and the statement has been changed.
    6. Comment: Special Condition No. 26. Should be deleted and 
replaced with Sec. 23.607, Amendment 23-48.
    FAA Response: The FAA agrees and Special Condition No. 26 has been 
deleted. Later amendment levels are adequate for this airplane.
    7. Comment: Special Condition No. 30--Pressurization. Special 
Condition No. 30 addresses the altitude-time histories of the cabin 
altitude following system and/or structural failures. The language and 
requirements defined in Special Condition No. 30, paragraphs (a)(2), 
(b)(1), and (b)(2), are a carry-over of early part 25 executive 
transport airplane special conditions developed for high altitude 
operation (above 40,000 feet). As discussed in the Federal Register, 
Volume 61, No. 109, dated June 5, 1996, part 25 special conditions were 
developed to address the consequences of decompression of executive 
transport airplanes operation at high altitudes. These early special 
conditions revised the requirements of Sec. 25.365, Pressurized Cabin 
Loads, Sec. 25.841, Pressurized Cabins, and Sec. 25.1447, Equipment 
Standards for Oxygen Dispensing Equipment and were intended to provide 
an evaluation of the consequences of cabin depressurization due to 
system and/or structural failures.
    However, the wording provided in Special Condition No. 30 is based 
on an earlier amendment (before Amendment 25-45) of Sec. 25.571, which 
allowed a choice between safe-life and fail-safe substantiation for 
airplane primary structure. The airplane inspections defined for 
Sec. 25.571 before Amendment 25-45 were not specifically based on crack 
growth for spectrum loading. Therefore, the executive transport 
airplane special conditions for operation at high altitudes specified a 
somewhat arbitrary criteria of structural failure considerations for a 
decompression event. Subsequent to the initial development of these 
executive transport high altitude special conditions, Sec. 25.571 was 
amended by Amendments 25-45 (1978) and 25-52 (1980) to require a damage 
tolerance evaluation of the airplane primary structure. The damage 
tolerance evaluation requires the development of inspection intervals 
and procedures for the detection of crack lengths associated with the 
decompression of critical vent areas. Since the structural failures to 
be considered for the decompression event are defined by the damage 
tolerance evaluation, the language shown in Special Condition No. 30, 
paragraphs (a)(2), (b)(1), and (b)(2), is not part of the current part 
25 regulatory requirements for High Altitude Operation of Subsonic 
Transport Airplanes.
    The commenter believes that the structural failures to be 
considered of a decompression event should be defined by the damage 
tolerance evaluation of the SJ30-2 airplane pressure vessel required by 
Special Condition No. 25, Pressure Vessel Integrity, and not by the 
predefined conditions outlined in Special Condition No. 30, paragraphs 
(a)(2), (b)(1), and (b)(2). Therefore, the commenter suggests their 
words, which reflect the more recent structural approach.
    FAA Response: The FAA agrees with the commenter and Special 
Condition No. 30 will be replaced.
    8. Comment: Special Condition No. 37, Operating Limitations. 
Paragraph (a)(3) change read ``V<INF>O</INF>'' to ``V<INF>A</INF>''.
    FAA Response: The FAA does not agree. V<INF>A</INF> was correctly 
changed to V<INF>O</INF> in an earlier part 23 amendment so it will 
remain unchanged in these special conditions.

Conclusion

    In view of the design features discussed for the SJ30-2 Model 
airplane, the following special conditions are issued to provide a 
level of safety equivalent to current business jets certificated to 
transport standards and expected by the user of this class of aircraft. 
This action is not a rule of general applicability and affects only the 
model/series of airplane identified in these final special conditions.

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 
101; and 14 CFR 11.28 and 11.49.

Adoption of Special Conditions

    Accordingly, pursuant to the authority delegated to me by the 
Administrator, the Federal Aviation Administration issues the following 
special conditions as part of the type certification basis for the Sino 
Swearingen Model SJ30-2 airplane:

1. Protection of Electrical and Electronic Systems From High 
Intensity Radiated Field

    Each system that performs critical functions must be designed and 
installed to ensure that the operations, and operational capabilities 
of these systems to perform critical functions, are not adversely 
affected when the airplane is exposed to high intensity radiated 
electromagnetic fields external to the airplane.
    For the purpose of these special conditions, the following 
definition applies:
    Critical Functions: Functions whose failure would contribute to, or 
cause, a failure condition that would prevent the continued safe flight 
and landing of the airplane.

2. Performance: General

    In addition to the requirements of Sec. 23.45, the following apply:
    (a) Unless otherwise prescribed, the applicant must select the 
takeoff, enroute, approach, and landing configurations for the 
airplane.
    (b) The airplane configurations may vary with weight, altitude, and 
temperature, to the extent that they are compatible with the operating 
procedures required by paragraph (c) of this special condition.
    (c) Unless otherwise prescribed, in determining the accelerate-stop 
distances, takeoff flight paths, takeoff distances, and landing 
distances, changes in the airplane's configuration, speed, power, and 
thrust, must be made in accordance with procedures established by the 
applicant for operation in service.
    (d) Procedures for the execution of balked landings and 
discontinued approaches associated with the conditions prescribed in 
special condition 10, paragraph (d), and special condition 12 must be 
established.
    (e) The procedures established under paragraphs (c) and (d) of this 
special condition must:
    (1) Be able to be consistently executed in service by crews of 
average skill;
    (2) Use methods or devices that are safe and reliable; and
    (3) Include allowance for any time delays, in the execution of the 
procedures, that may reasonably be expected in service.

3. Takeoff

    Instead of complying with Sec. 23.53, the following apply:
    (a) In special conditions 4, 5, 6, and 7, the takeoff speeds, the 
accelerate-stop distance, the takeoff path, the takeoff distance, and 
takeoff run described must be determined:

[[Page 58880]]

    (1) At each weight, altitude, and ambient temperature within the 
operation limits selected by the applicant; and
    (2) In the selected configuration for takeoff.
    (b) No takeoff made to determine the data required by this section 
may require exceptional piloting skill or alertness.
    (c) The takeoff data must be based on a smooth, dry, hard-surfaced 
runway.
    (d) The takeoff data must include, within the established 
operational limits of the airplane, the following operational 
correction factors:
    (1) Not more than 50 percent of nominal wind components along the 
takeoff path opposite to the direction of takeoff, and not less than 
150 percent of nominal wind components along the takeoff path in the 
direction of takeoff.
    (2) Effective runway gradients.

4. Takeoff Speeds

    Instead of compliance with Sec. 23.51, the following apply:
    (a) V<INF>1</INF> must be established in relation to 
V<INF>EF</INF>, as follows:
    (1) V<INF>EF</INF> is the calibrated airspeed at which the critical 
engine is assumed to fail. V<INF>EF</INF> must be selected by the 
applicant, but may not be less than V<INF>MCG</INF> determined under 
Sec. 23.149(f).
    (2) V<INF>1</INF>, in terms of calibrated airspeed, is the takeoff 
decision speed selected by the applicant; however, V<INF>1</INF> may 
not be less than V<INF>EF</INF> plus the speed gained with the critical 
engine inoperative during the time interval between the instant at 
which the critical engine failed and the instant at which the pilot 
recognizes and reacts to the engine failure, as indicated by the 
pilot's application of the first retarding means during the accelerate-
stop test.
    (b) V<INF>2 min</INF>, in terms of calibrated airspeed, may not be 
less than the following:
    (1) 1.2 V<INF>S1</INF>
    (2) 1.10 times V<INF>MC</INF> established under Sec. 23.149.
    (c) V<INF>2</INF>, in terms of calibrated airspeed, must be 
selected by the applicant to provide at least the gradient of climb 
required by special condition 10, paragraph (b), but may not be less 
than the following:
    (1) V<INF>2 min</INF>, and
    (2) V<INF>R</INF> plus the speed increment attained (in accordance 
with special condition 6, paragraph (c)(2)) before reaching a height of 
35 feet above the takeoff surface.
    (d) V<INF>MU</INF> is the calibrated airspeed at and above which 
the airplane can safely lift off the ground and continue the takeoff. 
V<INF>MU</INF> speeds must be selected by the applicant throughout the 
range of thrust-to-weight ratios to be certified. These speeds may be 
established from free-air data if these data are verified by ground 
takeoff tests.
    (e) V<INF>R</INF>, in terms of calibrated airspeed, must be 
selected in accordance with the following conditions of paragraphs 
(e)(1) through (e)(4) of this special condition:
    (1) V<INF>R</INF> may not be less than the following:
    (i) V<INF>1</INF>;
    (ii) 105 percent of V<INF>MC</INF>;
    (iii) The speed (determined in accordance with special condition 6, 
paragraph (c)(2)) that allows reaching V<INF>2</INF> before reaching a 
height of 35 feet above the takeoff surface; or
    (iv) A speed that, if the airplane is rotated at its maximum 
practicable rate, will result in a V<INF>LOF</INF> of not less than 110 
percent of V<INF>MU</INF> in the all-engines-operating condition and 
not less than 105 percent of V<INF>MU</INF> determined at the thrust-
to-weight ratio corresponding to the one-engine-inoperative condition.
    (2) For any given set of conditions (such as weight, configuration, 
and temperature), a single value of V<INF>R</INF>, obtained in 
accordance with this special condition, must be used to show compliance 
with both the one-engine-inoperative and the all-engines-operating 
takeoff provisions.
    (3) It must be shown that the one-engine-inoperative takeoff 
distance, using a rotation speed of 5 knots less than V<INF>R</INF>, 
established in accordance with paragraphs (e)(1) and (e)(2) of this 
special condition, does not exceed the corresponding one-engine-
inoperative takeoff distance using the established V<INF>R</INF>. The 
takeoff distances must be determined in accordance with special 
condition 7, paragraph (a)(1).
    (4) Reasonably expected variations in service from the established 
takeoff procedures for the operation of the airplane (such as over-
rotation of the airplane and out-of-trim conditions) may not result in 
unsafe flight characteristics or in marked increases in the scheduled 
takeoff distances established in accordance with special condition 7.
    (f) V<INF>LOF</INF> is the calibrated airspeed at which the 
airplane first becomes airborne.

5. Accelerate-Stop Distance

    In the absence of specific accelerate-stop distance requirements, 
the following apply:
    (a) The accelerate-stop distance is the sum of the distances 
necessary to--
    (1) Accelerate the airplane from a standing start to V<INF>EF</INF> 
with all engines operating;
    (2) Accelerate the airplane from V<INF>EF</INF> to V<INF>1</INF>, 
assuming that the critical engine fails at V<INF>EF</INF>; and
    (3) Come to a full stop from the point at which V<INF>1</INF> is 
reached assuming that, in the case of engine failure, the pilot has 
decided to stop as indicated by application of the first retarding 
means at the speed V<INF>1</INF>.
    (b) Means other than wheel brakes may be used to determine the 
accelerate-stop distance if that means--
    (1) Is safe and reliable;
    (2) Is used so that consistent results can be expected under normal 
operating conditions; and
    (3) Is such that exceptional skill is not required to control the 
airplane.
    (c) The landing gear must remain extended throughout the 
accelerate-stop distance.

6. Takeoff Path

    In the absence of specific takeoff path requirements, the following 
apply:
    (a) The takeoff path extends from a standing start to a point in 
the takeoff at which the airplane is 1,500 feet above the takeoff 
surface, or at which the transition from the takeoff to the enroute 
configuration is completed and a speed is reached at which compliance 
with special condition 10, paragraph (c), is shown, whichever point is 
higher. In addition, the following apply:
    (1) The takeoff path must be based on procedures prescribed in 
special condition 2.
    (2) The airplane must be accelerated on the ground to 
V<INF>EF</INF>, at which point the critical engine must be made 
inoperative and remain inoperative for the rest of the takeoff; and
    (3) After reaching V<INF>EF</INF>, the airplane must be accelerated 
to V<INF>2</INF>.
    (b) During the acceleration to speed V<INF>2</INF>, the nose gear 
may be raised off the ground at a speed not less than V<INF>R</INF>. 
However, landing gear retraction may not begin until the airplane is 
airborne.
    (c) During the takeoff path determination, in accordance with 
paragraphs (a) and (b) of this special condition, the following apply:
    (1) The slope of the airborne part of the takeoff path must be 
positive at each point;
    (2) The airplane must reach V<INF>2</INF> before it is 35 feet 
above the takeoff surface and must continue at a speed as close as 
practical to, but not less than, V<INF>2</INF> until it is 400 feet 
above the takeoff surface;
    (3) At each point along the takeoff path, starting at the point at 
which the airplane reaches 400 feet above the takeoff surface, the 
available gradient of climb may not be less than 1.2 percent;
    (4) Except for gear retraction, the airplane configuration may not 
be

[[Page 58881]]

changed, and no change in power or thrust that requires action by the 
pilot may be made, until the airplane is 400 feet above the takeoff 
surface.
    (d) The takeoff path must be determined by a continuous 
demonstrated takeoff or by synthesis from segments. If the takeoff path 
is determined by the segmental method, the following apply:
    (1) The segments must be clearly defined and must be related to the 
distinct changes in the configuration, speed, and power or thrust;
    (2) The weight of the airplane, the configuration, and the power or 
thrust must be constant throughout each segment and must correspond to 
the most critical condition prevailing in the segment;
    (3) The flight path must be based on the airplane's performance 
without ground effect; and
    (4) The takeoff path data must be checked by continuous 
demonstrated takeoffs, up to the point at which the airplane is out of 
ground effect and its speed is stabilized, to ensure that the path is 
conservative relative to the continuous path.

    Note: The airplane is considered to be out of the ground effect 
when it reaches a height equal to its wing span.

7. Takeoff Distance and Takeoff Run

    In the absence of specific takeoff distance and takeoff run 
requirements, the following apply:
    (a) Takeoff distance is the greater of the following:
    (1) The horizontal distance along the takeoff path from the start 
of the takeoff to the point at which the airplane is 35 feet above the 
takeoff surface, determined under special condition 6; or
    (2) 115 percent of the horizontal distance along the takeoff path, 
with all engines operating, from the start of the takeoff to the point 
at which the airplane is 35 feet above the takeoff surface, as 
determined by a procedure consistent with special condition 6.
    (b) If the takeoff distance includes a clear way, the takeoff run 
is the greater of:
    (1) The horizontal distance along the takeoff path from the start 
of the takeoff to a point equidistant between the point at which 
V<INF>LOF</INF> is reached and the point at which the airplane is 35 
feet above the takeoff surface, as determined under special condition 
6; or
    (2) 115 percent of the horizontal distance along the takeoff path, 
with all engines operating, from the start of the takeoff to a point 
equidistant between the point at which V<INF>LOF</INF> is reached and 
the point at which the airplane is 35 feet above the takeoff surface, 
determined by a procedure consistent with special condition 6.

8. Takeoff Flight Path

    In the absence of specific takeoff flight path requirements, the 
following apply:
    (a) The takeoff flight path begins 35 feet above the takeoff 
surface at the end of the takeoff distance determined in accordance 
with special condition 7.
    (b) The net takeoff flight path data must be determined so that 
they represent the actual takeoff flight paths (determined in 
accordance with special condition 6 and with paragraph (a) of this 
special condition) reduced at each point by a gradient of climb equal 
to 0.8 percent.
    (c) The prescribed reduction in climb gradient may be applied as an 
equivalent reduction in acceleration along that part of the takeoff 
flight path at which the airplane is accelerated in level flight.

9. Climb: General

    Instead of compliance with Sec. 23.63, the following applies: 
Compliance with the requirements of special conditions 10 and 12 must 
be shown at each weight, altitude, and ambient temperature within the 
operational limits established for the airplane and with the most 
unfavorable center of gravity for each configuration.

10. Climb: One Engine Inoperative

    Instead of compliance with Sec. 23.67, the following apply:
    (a) Takeoff; landing gear extended. In the critical takeoff 
configuration existing along the flight path (between the points at 
which the airplane reaches V<INF>LOF</INF> and at which the landing 
gear is fully retracted) and in the configuration used in special 
condition 6 without ground effect, unless there is a more critical 
power operating condition existing later along the flight path before 
the point at which the landing gear is fully retracted, the steady 
gradient of climb must be positive at V<INF>LOF</INF> and with the 
following:
    (1) The critical engine inoperative and the remaining engines at 
the power or thrust available when retraction of the landing gear 
begins in accordance with special condition 6, and
    (2) The weight equal to the weight existing when retraction of the 
landing gear begins, determined under special condition 6.
    (b) Takeoff; landing gear retracted. In the takeoff configuration 
existing at the point of the flight path at which the landing gear is 
fully retracted and in the configuration used in special condition 6, 
without ground effect, the steady gradient of climb may not be less 
than 2.4 percent at V<INF>2</INF> and with the following:
    (1) The critical engine inoperative, the remaining engines at the 
takeoff power or thrust available at the time the landing gear is fully 
retracted, determined under special condition 6 unless there is a more 
critical power operating condition existing later along the flight path 
but before the point where the airplane reaches a height of 400 feet 
above the takeoff surface; and
    (2) The weight equal to the weight existing when the airplane's 
landing gear is fully retracted, determined under special condition 6.
    (c) Final takeoff. In the enroute configuration at the end of the 
takeoff path, determined in accordance with special condition 6, the 
steady gradient of climb may not be less than 1.2 percent at not less 
than 1.25 V<INF>S</INF> and with the following:
    (1) The critical engine inoperative and the remaining engines at 
the available maximum continuous power or thrust; and
    (2) The weight equal to the weight existing at the end of the 
takeoff path, determined under special condition 6.
    (d) Approach. In the approach configuration corresponding to the 
normal all-engines-operating procedure in which V<INF>S</INF> for this 
configuration does not exceed 110 percent of the V<INF>S</INF> for the 
related landing configuration, the steady gradient of climb may not be 
less than 2.1 percent with the following:
    (1) The critical engine inoperative, the remaining engine at the 
available in-flight takeoff power or thrust;
    (2) The maximum landing weight; and
    (3) A climb speed established in connection with normal landing 
procedures, but not exceeding 1.5 V<INF>S</INF>.

11. Landing

    Instead of compliance with Sec. 23.75, the following apply:
    (a) The horizontal distance necessary to land and to come to a 
complete stop from a point 50 feet above the landing surface must be 
determined (for each weight, altitude, temperature, and wind within the 
operational limits established by the applicant for the airplane), as 
follows:
    (1) The airplane must be in the landing configuration.
    (2) A steady approach at a gradient of descent not greater than 5.2 
percent (3 degrees), with an airspeed of not less than V<INF>REF</INF>, 
determined in accordance with Sec. 23.73(b), must be maintained down to 
the 50-foot height.
    (3) Changes in configuration, power or thrust, and speed, must be 
made in accordance with the established procedures for service 
operation.

[[Page 58882]]

    (4) The landing must be made without excessive vertical 
acceleration, tendency to bounce, nose over, ground loop, or porpoise.
    (5) The landings may not require exceptional piloting skill or 
alertness.
    (6) It must be shown that a safe transition to the balked landing 
conditions of special condition 12 can be made from the conditions that 
exist at the 50-foot height.
    (b) The landing distance must be determined on a level, smooth, 
dry, hard-surfaced runway. In addition, the following apply:
    (1) The brakes may not be used so as to cause excessive wear of 
brakes or tires; and
    (2) Means other than wheel brakes may be used if that means is as 
follows:
    (i) Is safe and reliable;
    (ii) Is used so that consistent results can be expected in service; 
and
    (iii) Is such that exceptional skill is not required to control the 
airplane.
    (c) The landing distance data must include correction factors for 
not more than 50 percent of the nominal wind components along the 
landing path opposite to the direction of landing and not less than 150 
percent of the nominal wind components along the landing path in the 
direction of landing.
    (d) If any device is used that depends on the operation of any 
engine, and if the landing distance would be noticeably increased when 
a landing is made with that engine inoperative, the landing distance 
must be determined with that engine inoperative unless the use of 
compensating means will result in a landing distance not more than that 
with each engine operating.

12. Balked Landing

    Instead of compliance with Sec. 23.77, the following apply:
    In the landing configuration, the steady gradient of climb may not 
be less than 3.2 percent with the following:
    (a) The engines at the power or thrust that is available eight 
seconds after initiation of movement of the power or thrust controls 
from the minimum flight idle to the inflight takeoff position; and
    (b) A climb speed of not more than V<INF>REF</INF>, as defined in 
Sec. 23.73(a).

13. Stall Speed

    Instead of compliance with Sec. 23.49, the following apply:
    (a) V<INF>S</INF> is the calibrated stalling speed, or the minimum 
steady flight speed, in knots, at which the airplane is controllable 
with--
    (1) Zero thrust at the stalling speed, or, if the resultant thrust 
has no appreciable effect on the stalling speed, with engines idling 
and throttles closed;
    (2) The weight used when V<INF>S</INF> is being used as a factor to 
determine compliance with a required performance standard; and
    (3) The most unfavorable center of gravity allowable.
    (b) The stalling speed V<INF>S</INF> is the minimum speed obtained 
as follows:
    (1) Trim the airplane for straight flight at any speed not less 
than 1.2 V<INF>S</INF> or more than 1.4 V<INF>S</INF>. At a speed 
sufficiently above the stall speed to ensure steady conditions, apply 
the elevator control at a rate so that the airplane speed reduction 
does not exceed one knot per second.
    (2) Meet the flight characteristics provisions of special condition 
19.

14. Trim

    Instead of compliance with Sec. 23.161, the following apply:
    (a) General. Each airplane must meet the trim requirements of this 
special condition after being trimmed, and without further pressure 
upon or movement of the primary controls or their corresponding trim 
controls by the pilot or the automatic pilot.
    (b) Lateral and directional trim. The airplane must maintain 
lateral and directional trim with the most adverse lateral displacement 
of the center of gravity within the relevant operating limitations 
during normally expected conditions of operation (including operation 
at any speed from 1.4 V<INF>S1</INF> to V<INF>MO</INF>/M<INF>MO</INF>.)
    (c) Longitudinal trim. The airplane must maintain longitudinal trim 
during the following:
    (1) A climb with maximum continuous power at a speed not more than 
1.4 V<INF>S1</INF>, with the landing gear retracted, and the flaps in 
the following positions:
    (i) Retracted, and
    (ii) In the takeoff position.
    (2) A power approach with a 3 degree angle of descent, the landing 
gear extended, and with the following:
    (i) The wing flaps retracted and at a speed of 1.4 V<INF>S1</INF>; 
and
    (ii) The applicable airspeed and flap position used in showing 
compliance with special condition 11.
    (3) Level flight at any speed from 1.4 V<INF>S1</INF> to 
V<INF>MO</INF>/M<INF>MO</INF> with the landing gear and flaps 
retracted, and from 1.4 V<INF>S1</INF> to V<INF>LE</INF> with the 
landing gear extended.
    (d) Longitudinal, directional, and lateral trim. The airplane must 
maintain longitudinal, directional, and lateral trim (for the lateral 
trim, the angle of bank may not exceed five degrees) at 1.4 
V<INF>S1</INF> during climbing flight with the following:
    (1) The critical engine inoperative;
    (2) The remaining engine at maximum continuous power or thrust; and
    (3) The landing gear and flaps retracted.

15. Static Longitudinal Stability

    Instead of compliance with Sec. 23.173, the following apply:
    Under the conditions specified in special condition 16, the 
characteristics of the elevator control forces (including friction) 
must be as follows:
    (a) A pull must be required to obtain and maintain speeds below the 
specified trim speed, and a push must be required to obtain and 
maintain speeds above the specified trim speed. This must be shown at 
any speed that can be obtained except speeds higher than the landing 
gear or wing flap operating limit speeds or V<INF>FC</INF>/
M<INF>FC</INF>, whichever is appropriate, or lower than the minimum 
speed for steady unstalled flight.
    (b) The airspeed must return to within 10 percent of the original 
trim speed for the climb, approach, and landing conditions specified in 
special condition 16, paragraphs (a), (c), and (d), and must return to 
within 7.5 percent of the original trim speed for the cruising 
condition specified in special condition 16, paragraph (b), when the 
control force is slowly released from any speed within the range 
specified in paragraph (a) of this special condition.
    (c) The average gradient of the stable slope of the stick force 
versus speed curve may not be less than 1 pound for each 6 knots.
    (d) Within the free return speed range specified in paragraph (b) 
of this special condition, it is permissible for the airplane, without 
control forces, to stabilize on speeds above or below the desired trim 
speeds if exceptional attention on the part of the pilot is not 
required to return to and maintain the desired trim speed and altitude.

16. Demonstration of Static Longitudinal Stability

    Instead of compliance with Sec. 23.175, static longitudinal 
stability must be shown as follows:
    (a) Climb. The stick force curve must have a stable slope at speeds 
between 85 and 115 percent of the speed at which the airplane--
    (1) Is trimmed, with--
    (i) Wing flaps retracted;
    (ii) Landing gear retracted;
    (iii) Maximum takeoff weight; and
    (iv) The maximum power or thrust selected by the applicant as an 
operating limitation for use during climb; and
    (2) Is trimmed at the speed for best rate of climb except that the 
speed need not be less than 1.4 V<INF>S1</INF>.
    (b) Cruise. Static longitudinal stability must be shown in the 
cruise condition as follows:

[[Page 58883]]

    (1) With the landing gear retracted at high speed, the stick force 
curve must have a stable slope at all speeds within a range which is 
the greater of 15 percent of the trim speed plus the resulting free 
return speed range, or 50 knots plus the resulting free return speed 
range, above and below the trim speed (except that the speed range need 
not include speeds less than 1.4 V<INF>S1</INF>, nor speeds greater 
than V<INF>FC</INF>/M<INF>FC</INF>, nor speeds that require a stick 
force of more than 50 pounds), with--
    (i) The wing flaps retracted;
    (ii) The center of gravity in the most adverse position;
    (iii) The most critical weight between the maximum takeoff and 
maximum landing weights;
    (iv) The maximum cruising power selected by the applicant as an 
operating limitation, except that the power need not exceed that 
required at V<INF>MO</INF>/M<INF>MO</INF>; and
    (v) The airplane trimmed for level flight with the power required 
in paragraph (b)(1)(iv) of this special condition.
    (2) With the landing gear retracted at low speed, the stick force 
curve must have a stable slope at all speeds within a range which is 
the greater of 15 percent of the trim speed plus the resulting free 
return speed range, or 50 knots plus the resulting free return speed 
range, above and below the trim speed (except that the speed range need 
not include speeds less than 1.4 V<INF>S1</INF>, nor speeds greater 
than the minimum speed of the applicable speed range prescribed in 
paragraph (b)(1), nor speeds that require a stick force of more than 50 
pounds), with--
    (i) Wing flaps, center of gravity position, and weight as specified 
in paragraph (b)(1) of this special condition;
    (ii) Power required for level flight at a speed equal to 
(V<INF>MO</INF> + 1.4 V<INF>S1</INF>)/ 2; and
    (iii) The airplane trimmed for level flight with the power required 
in paragraph (b)(2)(ii) of this special condition.
    (3) With the landing gear extended, the stick force curve must have 
a stable slope at all speeds within a range which is the greater of 15 
percent of the trim speed plus the resulting free return speed range, 
or 50 knots plus the resulting free return speed range, above and below 
the trim speed (except that the speed range need not include speeds 
less than 1.4 V<INF>S1</INF>, nor speeds greater than V<INF>LE</INF>, 
nor speeds that require a stick force of more than 50 pounds), with--
    (i) Wing flap, center of gravity position, and weight as specified 
in paragraph (b)(1) of this section;
    (ii) The maximum cruising power selected by the applicant as an 
operating limitation, except that the power need not exceed that 
required for level flight at V<INF>LE</INF>; and
    (iii) The aircraft trimmed for level flight with the power required 
in paragraph (b)(3)(ii) of this section.
    (c) Approach. The stick force curve must have a stable slope at 
speeds between 1.1 V<INF>S1</INF> and 1.8 V<INF>S1</INF>, with--
    (1) Wing flaps in the approach position;
    (2) Landing gear retracted;
    (3) Maximum landing weight; and
    (4) The airplane trimmed at 1.4 V<INF>S1</INF> with enough power to 
maintain level flight at this speed.
    (d) Landing. The stick force curve must have a stable slope, and 
the stick force may not exceed 80 pounds, at speeds between 1.1 
V<INF>S0</INF> and 1.3 V<INF>S0</INF> with--
    (1) Wing flaps in the landing position;
    (2) Landing gear extended;
    (3) Maximum landing weight;
    (4) Power or thrust off on the engines; and
    (5) The airplane trimmed at 1.4 V<INF>S0</INF> with power or thrust 
off.

17. Static Directional and Lateral Stability

    Instead of compliance with Sec. 23.177, the following apply:
    (a) The static directional stability (as shown by the tendency to 
recover from a skid with the rudder free) must be positive for any 
landing gear and flap position, and it must be positive for any 
symmetrical power condition to speeds from 1.2 V<INF>S1</INF> up to 
V<INF>FE</INF>, V<INF>LE</INF>, or V<INF>FC</INF>/M<INF>FC</INF> (as 
appropriate).
    (b) The static lateral stability (as shown by the tendency to raise 
the low wing in a sideslip with the aileron controls free and for any 
landing gear position and flap position, and for any symmetrical power 
conditions) may not be negative at any airspeed (except speeds higher 
than V<INF>FE</INF> or V<INF>LE</INF>, when appropriate) in the 
following airspeed ranges:
    (1) From 1.2 V<INF>S1</INF> to V<INF>MO</INF>/M<INF>MO</INF>.
    (2) From V<INF>MO</INF>/M<INF>MO</INF> to V<INF>FC</INF>/
M<INF>FC</INF>, unless the Administrator finds that the divergence is--
    (i) Gradual;
    (ii) Easily recognizable by the pilot; and
    (iii) Easily controllable by the pilot.
    (c) In straight, steady, sideslips (unaccelerated forward slips) 
the aileron and rudder control movement and forces must be 
substantially proportional to the angle of the sideslip. The factor of 
proportionality must lie 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 when a rudder pedal force of 180 pounds 
is obtained, the rudder pedal forces may not reverse and increased 
rudder deflection must produce increased angles of sideslip. Unless the 
airplane has a yaw indicator, there must be enough bank accompanying 
sideslipping to clearly indicate any departure from steady unyawed 
flight.

18. Stall Demonstration

    Instead of compliance with Sec. 23.201, the following apply:
    (a) Stalls must be shown in straight flight and in 30 degree banked 
turns with--
    (1) Power off; and
    (2) The power necessary to maintain level flight at 1.6 
V<INF>S1</INF> (where V<INF>S1</INF> corresponds to the stalling speed 
with flaps in the approach position, the landing gear retracted, and 
maximum landing weight).
    (b) In each condition required by paragraph (a) of this section, it 
must be possible to meet the applicable requirements of special 
condition 19 with--
    (1) Flaps, landing gear, and deceleration devices in any likely 
combination of positions approved for operation;
    (2) Representative weights within the range for which certification 
is requested;
    (3) The most adverse center of gravity for recovery; and
    (4) The airplane trimmed for straight flight at the speed 
prescribed in special condition 13.
    (c) The following procedures must be used to show compliance with 
special condition 19:
    (1) Starting at a speed sufficiently above the stalling speed to 
ensure that a steady rate of speed reduction can be established, apply 
the longitudinal control so that the speed reduction does not exceed 
one knot per second until the airplane is stalled.
    (2) In addition, for turning flight stalls, apply the longitudinal 
control to achieve airspeed deceleration rates up to 3 knots per 
second.
    (3) As soon as the airplane is stalled, recover by normal recovery 
techniques.
    (d) The airplane is considered stalled when the behavior of the 
airplane gives the pilot a clear and distinctive indication of an 
acceptable nature that the airplane is stalled. Acceptable indications 
of a stall, occurring either individually or in combination, are--
    (1) A nose-down pitch that cannot be readily arrested;

[[Page 58884]]

    (2) Buffeting, of a magnitude and severity that is a strong and 
effective deterrent to further speed reduction; or
    (3) The pitch control reaches the aft stop and no further increase 
in pitch attitude occurs when the control is held full aft for a short 
time before recovery is initiated.

19. Stall Characteristics

    Instead of compliance with Sec. 23.203, the following applies:
    (a) It must be possible to produce and to correct roll and yaw by 
unreversed use of the aileron and rudder controls, up to the time the 
airplane is stalled. No abnormal nose up pitching may occur. The 
longitudinal control force must be positive up to and throughout the 
stall. In addition, it must be possible to promptly prevent stalling 
and to recover from a stall by normal use of the controls.
    (b) For level wing stalls, the roll occurring between the stall and 
the completion of the recovery may not exceed approximately 20 degrees.
    (c) For turning flight stalls, the action of the airplane after the 
stall may not be so violent or extreme as to make it difficult, with 
normal piloting skill, to effect a prompt recovery and to regain 
control of the airplane. The maximum bank angle that occurs during the 
recovery may not exceed--
    (1) Approximately 60 degrees in the original direction of the turn, 
or 30 degrees in the opposite direction, for deceleration rates up to 1 
knot per second; and
    (2) Approximately 90 degrees in the original direction of the turn, 
or 60 degrees in the opposite direction, for deceleration rates in 
excess of 1 knot per second.

20. Stall Warning

    Instead of compliance with Sec. 23.207, the following applies:
    (a) Stall warning with sufficient margin to prevent inadvertent 
stalling with the flaps and landing gear in any normal position must be 
clear and distinctive to the pilot in straight and turning flight.
    (b) The warning may be furnished either through the inherent 
aerodynamic qualities of the airplane or by a device that will give 
clearly distinguishable indications under expected conditions of 
flight. However, a visual stall warning device that requires the 
attention of the crew within the cockpit is not acceptable by itself. 
If a warning device is used, it must provide a warning in each of the 
airplane configurations prescribed in paragraph (a) of this special 
condition at the speed prescribed in paragraph (c) of this special 
condition.
    (c) The stall warning must begin at a speed exceeding the stalling 
speed (i.e., the speed at which the airplane stalls or the minimum 
speed demonstrated, whichever is applicable under the provisions of 
special condition 18, paragraph (d)) by seven percent or at any lesser 
margin if the stall warning has enough clarity, duration, 
distinctiveness, or similar properties.

21. Vibration and Buffeting

    Instead of compliance with Sec. 23.251, the following apply:
    (a) The airplane must be designed to withstand any vibration and 
buffeting that might occur in any likely operating condition. This must 
be shown by calculations, resonance tests, or other tests found 
necessary by the Administrator.
    (b) Each part of the airplane must be shown in flight to be free 
from excessive vibration, under any appropriate speed and power 
conditions up to V<INF>DF</INF>/M<INF>DF</INF>. The maximum speeds 
shown must be used in establishing the operating limitations of the 
airplane in accordance with special condition 34.
    (c) Except as provided in paragraph (d) of this special condition, 
there may be no buffeting condition in normal flight, including 
configuration changes during cruise, severe enough to interfere with 
the control of the airplane, to cause excessive fatigue to the 
flightcrew, or to cause structural damage. Stall warning buffeting 
within these limits is allowable.
    (d) There may be no perceptible buffeting condition in the cruise 
configuration in straight flight at any speed up to V<INF>MO</INF>/
M<INF>MO</INF>, except that stall warning buffeting is allowable.
    (e) With the airplane in the cruise configuration, the positive 
maneuvering load factors at which the onset of perceptible buffeting 
occurs must be determined for the ranges of airspeed or Mach Number, 
weight, and altitude for which the airplane is to be certified. The 
envelopes of load factor, speed, altitude, and weight must provide a 
sufficient range of speeds and load factors for normal operations. 
Probable inadvertent excursions beyond the boundaries of the buffet 
onset envelopes may not result in unsafe conditions.

22. High Speed Characteristics

    Instead of compliance with Sec. 23.253, the following apply:
    (a) Speed increase and recovery characteristics. The following 
speed increase and recovery characteristics must be met:
    (1) Operating conditions and characteristics likely to cause 
inadvertent speed increases (including upsets in pitch and roll) must 
be simulated with the airplane trimmed at any likely cruise speed up to 
V<INF>MO</INF>/M<INF>MO</INF>. These conditions and characteristics 
include gust upsets, inadvertent control movements, low stick force 
gradient in relation to control friction, passenger movement, leveling 
off from climb, and descent from Mach to airspeed limit altitudes.
    (2) Allowing for pilot reaction time after effective inherent or 
artificial speed warning occurs, it must be shown that the airplane can 
be recovered to a normal attitude and its speed reduced to 
V<INF>MO</INF>/M<INF>MO</INF> without the following:
    (i) Exceptional piloting strength or skill;
    (ii) Exceeding V<INF>D</INF>/M<INF>D</INF>, or V<INF>DF</INF>/
M<INF>DF</INF>, or the structural limitations; and
    (iii) Buffeting that would impair the pilot's ability to read the 
instruments or control the airplane for recovery.
    (3) There may be no control reversal about any axis at any speed up 
to V<INF>DF</INF>/M<INF>DF</INF> with the airplane trimmed at 
V<INF>MO</INF>/M<INF>MO</INF>. Any tendency of the airplane to pitch, 
roll, or yaw must be mild and readily controllable, using normal 
piloting techniques. When the airplane is trimmed at V<INF>MO</INF>/
M<INF>MO</INF>, the slope of the elevator control force versus speed 
curve need not be stable at speeds greater than V<INF>FC</INF>/
M<INF>FC</INF>, but there must be a push force at all speeds up to 
V<INF>DF</INF>/M<INF>DF</INF> and there must be no sudden or excessive 
reduction of elevator control force as V<INF>DF</INF>/M<INF>DF</INF> is 
reached.
    (b) Maximum speed for stability characteristics. V<INF>FC</INF>/
M<INF>FC</INF>. V<INF>FC</INF>/M<INF>FC</INF> is the maximum speed at 
which the requirements of special conditions 15, 16, 17, and 
Sec. 23.181 must be met with the flaps and landing gear retracted. It 
may not be less than a speed midway between V<INF>MO</INF>/
M<INF>MO</INF> and V<INF>DF</INF>/M<INF>DF</INF> except that, for 
altitudes where Mach number is the limiting factor, M<INF>FC</INF> need 
not exceed the Mach number at which effective speed warning occurs.

23. Flight Flutter Testing

    Instead of the term/speed V<INF>D</INF> in Sec. 23.629(b), use 
V<INF>DF</INF>/M<INF>DF</INF>.

24. Out-of-Trim Characteristics

    In the absence of specific requirements for out-of-trim 
characteristics, the following are applied:
    (a) From an initial condition with the airplane trimmed at cruise 
speeds up to V<INF>MO</INF>/M<INF>MO</INF>, the airplane must have 
satisfactory maneuvering stability and controllability with the degree 
of out-of-trim in both the airplane nose-up and

[[Page 58885]]

nose-down directions, which results from the greater of the following:
    (1) A three-second movement of the longitudinal trim system at its 
normal rate for the particular flight condition with no aerodynamic 
load (or an equivalent degree of trim for airplanes that do not have a 
power-operated trim system), except as limited by stops in the trim 
system including those required by Sec. 23.655(b) for adjustable 
stabilizers; or
    (2) The maximum mis-trim that can be sustained by the autopilot 
while maintaining level flight in the high speed cruising condition.
    (b) In the out-of-trim condition specified in paragraph (a) of this 
special condition, when the normal acceleration is varied from +1 g to 
the positive and negative values specified in paragraph (c) of this 
special condition, the following apply:
    (1) The stick force versus g curve must have a positive slope at 
any speed up to and including V<INF>FC</INF>/M<INF>FC</INF>; and
    (2) At speeds between V<INF>FC</INF>/M<INF>FC</INF> and 
V<INF>DF</INF>/M<INF>DF</INF>, the direction of the primary 
longitudinal control force may not reverse.
    (c) Except as provided in paragraphs (d) and (e) of this special 
condition, compliance with the provisions of paragraph (a) of this 
special condition must be demonstrated in flight over the acceleration 
range as follows:
    (1) -1 g to +2.5 g; or
    (2) 0 g to 2.0 g, and extrapolating by an acceptable method to -1 g 
and +2.5 g.
    (d) If the procedure set forth in paragraph (c)(2) of this special 
condition is used to demonstrate compliance and marginal conditions 
exist during flight test with regard to reversal of primary 
longitudinal control force, flight tests must be accomplished from the 
normal acceleration at which a marginal condition is found to exist to 
the applicable limit specified in paragraph (b)(1) of this special 
condition.
    (e) During flight tests required by paragraph (a) of this special 
condition, the limit maneuvering load factors, prescribed in 
Secs. 23.333(b) and 23.337, need not be exceeded. Also, the maneuvering 
load factors associated with probable inadvertent excursions beyond the 
boundaries of the buffet onset envelopes determined under special 
condition 21, paragraph (e), need not be exceeded. In addition, the 
entry speeds for flight test demonstrations at normal acceleration 
values less than 1 g must be limited to the extent necessary to 
accomplish a recovery without exceeding V<INF>DF</INF>/M<INF>DF</INF>.
    (f) In the out-of-trim condition specified in paragraph (a) of this 
special condition, it must be possible from an overspeed condition at 
V<INF>DF</INF>/M<INF>DF</INF> to produce at least 1.5 g for recovery by 
applying not more than 125 pounds of longitudinal control force using 
either the primary longitudinal control alone or the primary 
longitudinal control and the longitudinal trim system. If the 
longitudinal trim is used to assist in producing the required load 
factor, it must be shown at V<INF>DF</INF>/M<INF>DF</INF> that the 
longitudinal trim can be actuated in the airplane nose-up direction 
with the primary surface loaded to correspond to the least of the 
following airplane nose-up control forces:
    (1) The maximum control forces expected in service, as specified in 
Secs. 23.301 and 23.397.
    (2) The control force required to produce 1.5 g.
    (3) The control force corresponding to buffeting or other phenomena 
of such intensity that is a strong deterrent to further application of 
primary longitudinal control force.

25. Pressure Vessel Integrity

    (a) The maximum extent of failure and pressure vessel opening that 
can be demonstrated to comply with special condition 30 
(Pressurization) of these special conditions must be determined. It 
must be demonstrated by crack propagation and damage tolerance analysis 
supported by testing that a larger opening or a more severe failure 
than demonstrated will not occur in normal operations.
    (b) Inspection schedules and procedures must be established to 
ensure that cracks and normal fuselage leak rates will not deteriorate 
to the extent that an unsafe condition could exist during normal 
operation.
    (c) With regard to the fuselage structure design for cabin pressure 
capability above 45,000 feet, the pressure vessel structure, including 
doors and windows, must comply with Sec. 23.365(d), using a factor of 
1.67 instead of the 1.33 factor prescribed.

26. Fasteners

    This section has been deleted, current Sec. 23.607 is adequate.

27. Landing Gear

    The main landing gear system must be designed so that if it fails 
due to overloads during takeoff or landing (assuming the overloads to 
act in the upward and aft directions), the failure mode is not likely 
to cause the spillage of enough fuel from any fuel system in the 
fuselage to constitute a fire hazard.

28. Ventilation

    In addition to the requirements of Sec. 23.831(b), the ventilation 
system must be designed to provide a sufficient amount of 
uncontaminated air to enable the crewmembers to perform their duties 
without undue discomfort or fatigue and to provide reasonable passenger 
comfort during normal operating conditions and in the event of any 
probable failure of any system on the airplane that would adversely 
affect the cabin ventilating air. For normal operations, crewmembers 
and passengers must be provided with at least 10 cubic feet of fresh 
air per minute per person, or the equivalent in filtered recirculated 
air, based on the volume and composition at the corresponding cabin 
pressure altitude of no more than 8,000 feet.

29. Air Conditioning

    In addition to the requirements of Sec. 23.831, cabin cooling 
systems must be designed to meet the following conditions during flight 
above 15,000 feet MSL:
    (a) After any probable failure, the cabin temperature/time history 
may not exceed the values shown in Figure 1. During this time, the 
humidity shall never exceed a level that corresponds to a water vapor 
pressure of 20mm Hg. Time = 0 minutes when the flightcrew recognizes 
the failure.
    (b) After any improbable failure, the cabin temperature/time 
history may not exceed the values shown in Figure 2. During this time, 
the humidity shall never exceed a level that corresponds to a water 
vapor pressure of 20mm Hg. Time = 0 minutes when the flightcrew 
recognizes the failure.

30. Pressurization

    Instead of compliance with Sec. 23.841, the following apply:
    (a) Pressurized cabins must be equipped to provide a cabin pressure 
altitude of not more than 8,000 feet at the maximum operating altitude 
of the airplane under normal operating conditions.
    (1) If certification for operation above 25,000 feet is requested, 
the airplane must be designed so that occupants will not be exposed to 
cabin pressure altitudes in excess of 15,000 feet after any probable 
failure condition in the pressurization system.
    (2) The airplane must be designed so that occupants will not be 
exposed to a cabin pressure altitude that exceeds that following after 
decompression from any failure conditions not shown to be extremely 
improbable:
    (i) Twenty-five thousand (25,000) feet for more than 2 minutes; or

[[Page 58886]]

    (ii) Forty thousand (40,000) feet for any duration.
    (3) Fuselage structure, engine and system failures are to be 
considered in evaluating the cabin decompression.
    (b) Pressurized cabins must have at least the following valves, 
controls, and indicators for controlling cabin pressure:
    (1) Two pressure relief valves to automatically limit the positive 
pressure differential to a predetermined value at the maximum rate of 
flow delivered by the pressure source. The combined capacity of the 
relief valves must be large enough so that the failure of any one valve 
would not cause an appreciable rise in the pressure differential. The 
pressure differential is positive when the internal pressure is greater 
than the external.
    (2) Two reverse pressure differential relief valves (or their 
equivalents) to automatically prevent a negative pressure differential 
that would damage the structure. One valve is enough, however, if it is 
of a design that reasonably precludes its malfunctioning.
    (3) A means by which the pressure differential can be rapidly 
equalized.
    (4) An automatic or manual regulator for controlling the intake or 
exhaust airflow, or both, for maintaining the required internal 
pressure and airflow rates.
    (5) Instruments at the pilot station to show the pressure 
differential, the cabin pressure altitude, and the rate of change of 
the cabin pressure altitude.
    (6) Warning indication at the pilot station to indicate when the 
safe or preset pressure differential and cabin pressure altitude limits 
are exceeded. Appropriate warning marking on the cabin pressure 
differential indicator meets the warning requirement for pressure 
differential limits and an aural or visual signal (in addition to cabin 
altitude indicating means) meets the warning requirement for cabin 
pressure altitude limits if it warns the flight crew when the cabin 
pressure altitude exceeds 10,000 feet.
    (7) A warning placard at the pilot station, if the structure is not 
designed for pressure differentials up to the maximum relief valve 
setting in combination with landing loads.
    (8) The pressure sensors necessary to meet the requirements of 
paragraphs (b)(5) and (b)(6) of this section and Sec. 23.1447, 
paragraphs (e) and (f), must be located and the sensing system must be 
designed so that, in the event of low of cabin pressure, the warning 
and automatic presentation devices, required by those provisions, will 
be actuated without any delay that would significantly increase the 
hazards resulting from decompression.

31. Airspeed Indicating System

    In addition to the requirements of Sec. 23.1323, the following 
apply:
    (a) The airspeed indicating system must be calibrated to determine 
the system error in flight and during the accelerate-takeoff ground 
run. The ground run calibration must be determined as follows:
    (1) From 0.8 of the minimum value of V<INF>1</INF> to the maximum 
value of V<INF>2,</INF>, considering the approved ranges of altitude 
and weight; and
    (2) With the flaps and power settings corresponding to the values 
determined in the establishment of the takeoff path under special 
condition 6, assuming that the critical engine fails at the minimum 
value of V<INF>1</INF>.
    (b) The information showing the relationship between IAS and CAS, 
determined in accordance with paragraph (a) of this special condition, 
must be shown in the Airplane Flight Manual.

32. Static Pressure System

    In addition to the requirements of Sec. 23.1325, the following 
apply:
    (a) The altimeter system calibration required by Sec. 23.1325(e) 
must be shown in the Airplane Flight Manual.
    (b) If an altimeter system is fitted with a device that provides 
corrections to the altimeter indication, the device must be designed 
and installed in such manner that it can be by-passed when it 
malfunctions, unless an alternate altimeter system is provided. Each 
correction device must be fitted with a means for indicating the 
occurrence of reasonably probable malfunctions, including power 
failure, to the flightcrew. The indicating means must be effective for 
any cockpit lighting condition likely to occur.

33. Oxygen Equipment and Supply

    (a) In addition to the requirements of Sec. 23.1441(d), the 
following applies: A quick-donning oxygen mask system with a pressure-
demand, mask mounted regulator must be provided for the flightcrew. It 
must be shown that each quick-donning mask can, with one hand and 
within 5 seconds, be placed on the face from its ready position, 
properly secured, sealed, and supplying oxygen upon demand.
    (b) In addition to the requirements of Sec. 23.1443, the following 
applies: A continuous flow oxygen system must be provided for the 
passengers.
    (c) In addition to the requirements of Sec. 23.1445, the following 
applies: If the flightcrew and passengers share a common source of 
oxygen, a means to separately reserve the minimum supply required by 
the flightcrew must be provided.

34. Maximum Operating Limit Speed

    Instead of compliance with Sec. 23.1505(c), the following applies: 
The maximum operating limit speed (V<INF>MO</INF>/M<INF>MO</INF> 
airspeed or Mach number, whichever is critical at a particular 
altitude) is a speed that may not be deliberately exceeded in any 
regime of flight (climb, cruise, or descent), unless a higher speed is 
authorized for flight test or pilot training operations. 
V<INF>MO</INF>/M<INF>MO</INF> must be established so that it is not 
greater than the design cruising speed, V<INF>C</INF>, and so that it 
is sufficiently below V<INF>D</INF>/M<INF>D</INF>, or V<INF>DF</INF>/
M<INF>DF</INF>, to make it highly improbable that the latter speeds 
will be inadvertently exceeded in operations. The speed margin between 
V<INF>MO</INF>/M<INF>MO</INF> and V<INF>D</INF>/M<INF>D</INF>, or 
V<INF>DF</INF>/M<INF>DF</INF>, may not be less than that determined 
under Sec. 23.335(b) or found necessary during the flight tests 
conducted under special condition 22.

35. Minimum Flightcrew

    Instead of compliance with Sec. 23.1523, the following apply:
    The minimum flightcrew must be established so that it is sufficient 
for safe operation considering:
    (a) The workload on individual flightcrew members and each 
flightcrew member workload determination must consider the following:
    (1) Flight path control,
    (2) Collision avoidance,
    (3) Navigation,
    (4) Communications,
    (5) Operation and monitoring of all essential airplane systems,
    (6) Command decisions, and
    (7) The accessibility and ease of operation of necessary controls 
by the appropriate flightcrew member during all normal and emergency 
operations when at the flightcrew member station.
    (b) The accessibility and ease of operation of necessary controls 
by the appropriate flightcrew member; and
    (c) The kinds of operation authorized under Sec. 23.1525.

36. Airplane Flight Manual

    Instead of compliance with Sec. 23.1581, the following applies:
    (a) Furnishing information. An Airplane Flight Manual must be 
furnished with each airplane, and it must contain the following:
    (1) Information required by special conditions 37, 38, and 39.
    (2) Other information that is necessary for safe operation because 
of design, operating, or handling characteristics.

[[Page 58887]]

    (3) Any limitation, procedure, or other information established as 
a condition of compliance with the applicable noise standards of part 
36 of this chapter.
    (b) Approved Information. Each part of the manual listed in special 
conditions 37, 38, and 39, that is appropriate to the airplane, must be 
furnished, verified, and approved, and must be segregated, identified, 
and clearly distinguished from each unapproved part of that manual.
    (c) Airplane Flight Manual. Each Airplane Flight Manual must 
include a table of contents if the complexity of the manual indicates a 
need for it.
    (d) Airplane Flight Manual. Each page of the Airplane Flight Manual 
containing information prescribed in this section must be of a type 
that is not easily erased, disfigured, or misplaced, and is capable of 
being inserted in a manual provided by the applicant, or in a folder, 
or in any other permanent binder.
    (e) Airplane Flight Manual. Provision must be made for stowing the 
Airplane Flight Manual in a suitable fixed container that is readily 
accessible to the pilot.
    (f) Revisions and amendments. Each Airplane Flight Manual (AFM) 
must contain a means for recording the incorporation of revisions and 
amendments.

37. Operating Limitations

    Instead of the requirements of Sec. 23.1583, the following apply:
    (a) Airspeed limitations. The following airspeed limitations and 
any other airspeed limitations necessary for safe operation must be 
furnished:
    (1) The maximum operating limit speed, V<INF>MO</INF>/
M<INF>MO</INF>, and a statement that this speed limit may not be 
deliberately exceeded in any regime of flight (climb, cruise, or 
descent) unless a higher speed is authorized for flight test or pilot 
training.
    (2) If an airspeed limitation is based upon compressibility 
effects, a statement to this effect and information as to any symptoms, 
the probable behavior of the airplane, and the recommended recovery 
procedures.
    (3) The maneuvering speed, V<INF>O</INF>, and a statement that full 
application of rudder and aileron controls, as well as maneuvers that 
involve angles of attack near the stall, should be confined to speeds 
below this value.
    (4) The maximum speed for flap extension, V<INF>FE</INF>, for the 
takeoff, approach, and landing positions.
    (5) The landing gear operating speed or speeds, V<INF>LO</INF>.
    (6) The landing gear extended speed, V<INF>LE</INF> if greater than 
V<INF>LO</INF>, and a statement that this is the maximum speed at which 
the airplane can be safely flown with the landing gear extended.
    (b) Powerplant limitations. The following information must be 
furnished:
    (1) Limitations required by Sec. 23.1521.
    (2) Explanation of the limitations, when appropriate.
    (3) Information necessary for marking the instruments, required by 
Sec. 23.1549 through Sec. 23.1553.
    (c) Weight and loading distribution. The weight and extreme forward 
and aft center of gravity limits required by Secs. 23.23 and 23.25 must 
be furnished in the Airplane Flight Manual. In addition, all of the 
following information and the information required by Sec. 23.1589 must 
be presented either in the Airplane Flight Manual or in a separate 
weight and balance control and loading document, which is incorporated 
by reference in the Airplane Flight Manual:
    (1) The condition of the airplane and the items included in the 
empty weight, as defined in accordance with Sec. 23.29.
    (2) Loading instructions necessary to ensure loading of the 
airplane within the weight and center of gravity limits, and to 
maintain the loading within these limits in flight.
    (d) Maneuvers. A statement that acrobatic maneuvers, including 
spins, are not authorized.
    (e) Maneuvering flight load factors. The positive maneuvering limit 
load factors for which the structure is proven, described in terms of 
accelerations, and a statement that these accelerations limit the angle 
of bank in turns and limit the severity of pull-up maneuvers must be 
furnished.
    (f) Flightcrew. The number and functions of the minimum flightcrew 
determined under special condition 35 must be furnished.
    (g) Kinds of operation. The kinds of operation (such as VFR, IFR, 
day, or night) and the meteorological conditions in which the airplane 
may or may not be used must be furnished. Any installed equipment that 
affects any operating limitation must be listed and identified as to 
operational function.
    (h) Additional operating limitations must be established as 
follows: (1) The maximum takeoff weights must be established as the 
weights at which compliance is shown with the applicable provisions of 
part 23 (including the takeoff climb provisions of special condition 
10, paragraphs (a) through (c), for altitudes and ambient 
temperatures).
    (2) The maximum landing weights must be established as the weights 
at which compliance is shown with the applicable provisions of part 23 
(including the approach climb and balked landing climb provisions of 
special conditions 10, paragraph (d), and 12 for altitudes and ambient 
temperatures).
    (3) The minimum takeoff distances must be established as the 
distances at which compliance is shown with the applicable provisions 
of part 23 (including the provisions of special conditions 5 and 7 for 
weights, altitudes, temperatures, wind components, and runway 
gradients).
    (4) The extremes for variable factors (such as altitude, 
temperature, wind, and runway gradients) are those at which compliance 
with the applicable provision of part 23 and these special conditions 
is shown.
    (i) Maximum operating altitude. The maximum altitude established 
under Sec. 23.1527 must be furnished.
    (j) Maximum passenger seating configuration. The maximum passenger 
seating configuration must be furnished.

38. Operating Procedures

    Instead of the requirements of Sec. 23.1585, the following applies:
    (a) Information and instruction regarding the peculiarities of 
normal operations (including starting and warming the engines, taxiing, 
operation of wing flaps, slats, landing gear, speed brake, and the 
automatic pilot) must be furnished, together with recommended 
procedures for the following:
    (1) Engine failure (including minimum speeds, trim, operation of 
the remaining engine, and operation of flaps);
    (2) Restarting turbine engines in flight (including the effects of 
altitude);
    (3) Fire, decompression, and similar emergencies;
    (4) Use of ice protection equipment;
    (5) Operation in turbulence (including recommended turbulence 
penetration airspeeds, flight peculiarities, and special control 
instructions);
    (6) The demonstrated crosswind velocity and procedures and 
information pertinent to operation of the airplane in crosswinds.
    (b) Information identifying each operating condition in which the 
fuel system independence prescribed in Sec. 23.953 is necessary for 
safety must be furnished, together with instructions for placing the 
fuel system in a configuration used to show compliance with that 
section.
    (c) For each airplane showing compliance with Sec. 23.1353(g)(2) or 
(g)(3), the operating procedures for disconnecting the battery from its 
charging source must be furnished.
    (d) If the unusable fuel supply in any tank exceeds 5 percent of 
the tank

[[Page 58888]]

capacity, or 1 gallon, whichever is greater, information must be 
furnished indicating that, when the fuel quantity indicator reads 
``zero'' in level flight, any fuel remaining in the fuel tank cannot be 
used safely in flight.
    (e) Information on the total quantity of usable fuel for each fuel 
tank must be furnished.
    (f) The buffet onset envelopes determined under special condition 
21 must be furnished. The buffet onset envelopes presented may reflect 
the center of gravity at which the airplane is normally loaded during 
cruise if corrections for the effect of different center of gravity 
locations are furnished.

39. Performance Information

    Instead of the requirements of Sec. 23.1587, the following applies:
    (a) Each Airplane Flight Manual must contain information to permit 
conversion of the indicated temperature to free air temperature if 
other than a free air temperature indicator is used to comply with the 
requirements of Sec. 23.1303(d).
    (b) Each Airplane Flight Manual must contain the performance 
information computed under the applicable provisions of this part for 
the weights, altitudes, temperatures, wind components, and runway 
gradients, as applicable, within the operational limits of the 
airplane, and must contain the following:
    (1) The conditions under which the performance information was 
obtained, including the speeds associated with the performance 
information.
    (2) V<INF>S</INF> determined in accordance with special condition 
13.
    (3) The following performance information (determined by 
extrapolation and computed for the range of weights between the maximum 
landing and maximum takeoff weights):
    (i) Climb in the landing configuration.
    (ii) Climb in the approach configuration.
    (iii) Landing distance.
    (4) Procedures established under special condition 2, paragraphs 
(c), (d), and (e), that are related to the limitations and information 
required by paragraph (h) of special condition 37 and by this 
paragraph. These procedures must be in the form of guidance material, 
including any relevant limitations or information.
    (5) An explanation of significant or unusual flight or ground 
handling characteristics of the airplane.

    Issued in Kansas City, Missouri on October 15, 1997.
Mary Ellen A. Schutt,
Acting Manager, Small Airplane Directorate, Aircraft Certification 
Service.

BILLING CODE 4910-13-P

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[GRAPHIC] [TIFF OMITTED] TR31OC97.003



[[Page 58890]]

[GRAPHIC] [TIFF OMITTED] TR31OC97.004


[FR Doc. 97-28937 Filed 10-30-97; 8:45 am]
BILLING CODE 4910-13-C