[Federal Register Volume 62, Number 35 (Friday, February 21, 1997)]
[Proposed Rules]
[Pages 7950-7964]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 97-4353]


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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 23

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


Special Conditions; Sino Swearingen Model SJ30-2 Airplane

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Notice of proposed special conditions.

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SUMMARY: This notice proposes special conditions for the Sino 
Swearingen Aircraft Company Model SJ30-2 airplane. This new airplane 
will have novel and unusual design features not typically associated 
with normal, utility, acrobatic, and commuter category airplanes. These 
design features include a high operating altitude (49,000 feet), engine 
location, swept wings and stabilizer, performance characteristics, 
large fuel capacity, and protection for the electronic engine control 
system, flight, and navigation system from high intensity radiated 
fields, for which the applicable regulations do not contain adequate or 
appropriate airworthiness standards. This notice contains the 
additional airworthiness standards which the Administrator considers 
necessary to establish a level of safety equivalent to the 
airworthiness standards applicable to these airplanes.

DATES: Comments must be received on or before March 24, 1997.

ADDRESSES: Comments on this proposal may be mailed in duplicate to: 
Federal Aviation Administration, Office of the Assistant Chief Counsel, 
ACE-7, Attention: Rules Docket Clerk, Docket No. 135CE, Room No. 1558, 
601 East 12th Street, Kansas City, Missouri 64106. All comments must be 
marked: Docket No 135CE. Comments may be inspected in the Rules Docket 
weekdays, except Federal holidays, between 7:30 a.m. and 4 p.m.

FOR FURTHER INFORMATION CONTACT: 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:

Comments Invited

    Interested persons are invited to participate in the making of 
these special conditions by submitting such written data, views, or 
arguments as they may desire. Communications should identify the 
regulatory docket or notice number and be submitted in duplicate to the 
address specified above. All communications received on or before the 
closing date for comments specified above will be considered by the 
Administrator before taking further rulemaking action on this proposal. 
Commenters wishing the FAA to acknowledge receipt of their comments 
submitted in response to this notice must include a self-addressed, 
stamped postcard on which the following statement is made: ``Comments 
to Docket No. 135CE.'' The postcard will be date stamped and returned 
to the commenter. The proposals contained in this notice may be changed 
in light of the comments received. All comments received will be 
available, both before and after the closing date for comments, in the 
rules docket for examination by interested parties. A report 
summarizing each substantive public contact with FAA personnel 
concerned with this rulemaking will be filed in the docket.

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/M<INF>MO of 320 kts/M=.83, and has 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 23-51, effective 
March 11, 1996; 14 CFR Part 36, effective

[[Page 7951]]

December 1, 1969, through the amendment effective on the date of type 
certification; 14 CFR Part 34; exemptions, if any; and any special 
conditions that may result from this notice.

Discussion

    Sino Swearingen plans to incorporate certain novel and unusual 
design features into the SJ30-2 airplane for which the airworthiness 
regulations do not contain adequate or appropriate safety standards. 
These features include engine location, operation up to an altitude of 
49,000 feet, and certain performance characteristics necessary for this 
type of airplane that were not envisioned by the existing regulations.
    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)

    The aviation industry uses electrical and electronic systems that 
perform functions required for continued safe flight and landing. Due 
to the sensitive solid state components in analog and digital 
electronics circuits, these systems, if unprotected, are 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 electromagnetic environment has changed from the 
time when the current requirements were developed. Also, the population 
of transmitters has increased significantly and they are radiating 
higher energy levels. There is, however, uncertainty concerning the 
effectiveness of shielding for HIRF. Additionally, coupling to cockpit 
installed equipment through the cockpit window apertures is undefined.
    The combined effect of the technological advances in aircraft 
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 aircraft. 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 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
500-2000......................................           70           70
2-30 MHz......................................          200          200
30-70.........................................           30           30
70-100........................................           30           30
100-200.......................................          150           33
200-400.......................................           70           70
400-700.......................................         4020          935
700-1000......................................         1700          170
1-2 GHz.......................................         5000          990
2-4...........................................         6680          840
4-6...........................................         6850          310
6-8...........................................         3600          670
8-12..........................................         3500         1270
12-18.........................................         3500          360
18-40.........................................         2100          750
------------------------------------------------------------------------

or:
    (2) The applicant may demonstrate by a laboratory test that the 
electrical and electronic systems that perform critical functions can 
withstand a peak electromagnetic field strength of 100 volts per meter 
(v/m) peak electrical field strength, from 10KHz to 18GHz. When using a 
laboratory 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 aircraft. 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 the 
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 a combination of 
these. Service experience alone is not acceptable since such experience 
in 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 double-
slotted 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 significant 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

[[Page 7952]]

precise test measurements or specific detection systems. This phenomena 
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/N<INF>2 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, 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 are proposed for 
the performance requirements of takeoff, takeoff speeds, accelerate-
stop distance, takeoff path, takeoff distance, takeoff run, and takeoff 
flight path.
    Additionally, procedures for takeoff, accelerate stop, and landing 
are proposed as those established for operation in service and be 
executable by pilots of average skill and include reasonably expected 
time delays.

Climb

    To maintain a level of safety that is consistent with the 
requirements of the proposed special conditions for takeoff, takeoff 
speeds, takeoff path, takeoff distance, and takeoff run, it is 
appropriate to propose associate requirements that specify climb 
gradients, airplane configurations, and consideration of atmospheric 
conditions that will be encountered. Special conditions are proposed 
for climb with one engine inoperative, balked landing climb, and 
general climb conditions.

Landing

    Landing distance determined for the same parameters, plus the 
effects of wind, 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. Special conditions are also proposed to cover 
these items.

Trim

    Special conditions are issued to maintain a level of safety that is 
consistent with the use of V<INF>MO/M<INF>MO 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.

Demonstration of Static Longitudinal Stability

    To maintain a level of safety consistent with the proposed static 
longitudinal stability requirements, it is necessary to establish 
corresponding requirements for the demonstration of static longitudinal 
stability. Current standards in part 23 did not envision this type of 
airplane and the associated stability considerations proposed. In 
keeping with the concept of V<INF>MO/M<INF>MO 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/M<INF>MO of 14 CFR part 23 to the 
maximum speed for stability characteristics, V<INF>FC/M<INF>FC, for 
this airplane. A special condition to do this is proposed.

Static Directional and Lateral Stability

    In keeping with the concept of V<INF>MO/M<INF>MO 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/M<INF>MO of part 23 to the maximum 
speed for stability characteristics, V<INF>FC/M<INF>FC, for this 
airplane. A special condition to do this is proposed.
    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 flight control system 
reliability requirement are 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 requirement.

Stall Characteristics

    In order to maintain consistency with the level of safety 
previously applied to other jet powered small airplanes, it is 
appropriate to specify the conditions under which level flight, turning 
flight, and accelerated entry stall characteristics should be 
demonstrated. Current rules contained in part 23 did not envision this 
high performance airplane with the associated high thrust-to-weight 
ratio. Special conditions are required to define stall characteristics 
demonstrations.

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. A special condition is proposed that 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/M<INF>MO.

[[Page 7953]]

High Speed Characteristics and Maximum Operating Limit Speed

    The Sino Swearingen Model SJ30-2 will be operated at high altitude 
and high speeds and 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/M<INF>DF, the highest demonstrated 
flight speed for the type design, is appropriate for this airplane. 
This permits V<INF>D/M<INF>D the design dive speed, to be higher than 
the speed actually required to be demonstrated in flight. Accordingly, 
special conditions are proposed to allow determination of a maximum 
demonstrated flight speed and to relate the determination of V<INF>MO/
M<INF>MO to the speed V<INF>DF/M<INF>DF.

Flight Flutter Tests

    Flight flutter test special conditions are proposed to V<INF>DF/
M<INF>DF rather than to V<INF>D in keeping with the V<INF>DF/M<INF>DF 
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. Special conditions are proposed that 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. The proposal would 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/M<INF>DF.

Pressure Vessel Integrity

    Damage tolerance methods are proposed to be used to ensure pressure 
vessel integrity while operating at the higher altitudes instead of the 
1/2 bay crack criterion used in some previous special conditions. 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. Initial crack sizes for detection are 
determined under Sec. 23.573. The cabin altitude after failure must not 
exceed the cabin altitude/time curve limits shown in Figures 3 and 4.

Flight Control System Integrity

    The Sino Swearingen Model SJ30-2 will be operated at high altitude 
and speeds such that a reduction or loss of pitch, yaw, or roll control 
capability or response could preclude continued flight and landing 
within the design limitations of the airplane using normal pilot skill 
and strength. Consequently, a greater reliability of the fasteners in 
the flight control system is necessary than previously considered. 
Removable fasteners whose loss could result in the conditions described 
above are required to have dual locking devices.

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 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. A special condition is proposed based on 14 CFR part 25, 
Sec. 25.721(a)(1) that 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 conditions 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 applied to 
previous part 23 jet airplanes, 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. Current standards

[[Page 7954]]

in part 23 did not envision this type of airplane and the associated 
airspeed indicating system requirements. Special conditions are 
proposed to establish airspeed indicating system calibration and pitot 
tube ice protection requirements applicable to transport category jet 
airplanes.

Static Pressure System

    To maintain a level of safety consistent with that applied to 
previous part 23 jet airplanes, and to be consistent with the 
establishment of speed schedule performance requirements, it is 
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 are proposed to 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 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 are proposed to 
add maximum takeoff weights, maximum landing weights, and minimum 
takeoff distances as limitations in the AFM. Additionally, special 
conditions are proposed to add takeoff flight path and procedures 
necessary to achieve the performance in the limitations section as 
information in the AFM.

Conclusion

    In view of the design features discussed for the SJ30-2 Model 
airplane, the following special conditions are proposed. 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 11.29.

The Proposed Special Conditions

    Accordingly, pursuant to the authority delegated to me by the 
Administrator, the Federal Aviation Administration proposes 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 operation and operational capabilities of 
these systems to perform critical functions are not adversely affected 
when the airplane is exposed to high intensity radiated fields external 
to the airplane.

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 conditions 10(d) and 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:
    (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 must be established in relation to V<INF>EF, as 
follows:
    (1) V<INF>EF is the calibrated airspeed at which the critical 
engine is assumed to fail. V<INF>EF must be selected by the applicant, 
but may not be less than V<INF>MCG determined under Sec. 23.149(f).
    (2) V<INF>1, in terms of calibrated airspeed, is the takeoff 
decision speed selected by the applicant; however, V<INF>1 may not be 
less than V<INF>EF 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, in terms of calibrated airspeed, may not be less 
than the following:
    (1) 1.2 V<INF>S1
    (2) 1.10 times V<INF>MC established under Sec. 23.149.
    (c) V<INF>2, in terms of calibrated airspeed, must be selected by 
the applicant to provide at least the gradient of climb required by 
special condition 10,

[[Page 7955]]

paragraph (b), but may not be less than the following:
    (1) V<INF>2 min, and
    (2) V<INF>R plus the speed increment attained (in accordance with 
special condition 6(c)(2)) before reaching a height of 35 feet above 
the takeoff surface.
    (d) V<INF>MU is the calibrated airspeed at and above which the 
airplane can safely lift off the ground and continue the takeoff. 
V<INF>MU 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, 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 may not be less than the following:
    (i) V<INF>1;
    (ii) 105 percent of V<INF>MC;
    (iii) The speed (determined in accordance with special condition 6, 
paragraph (c)(2)) that allows reaching V<INF>2 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 of not less than 110 
percent of V<INF>MU in the all-engines-operating condition and not less 
than 105 percent of V<INF>MU 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, 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, 
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. The takeoff 
distances must be determined in accordance with special condition 7, 
paragraph (a)(1).
    (4) Reasonably expecting 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 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 with 
all engines operating;
    (2) Accelerate the airplane from V<INF>EF to V<INF>1, assuming that 
the critical engine fails at VEF; and
    (3) Come to a full stop from the point at which V<INF>1 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 V1.
    (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, 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, the airplane must be accelerated to 
V<INF>2.
    (b) During the acceleration to speed V<INF>2, the nose gear may be 
raised off the ground at a speed not less than VR. 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 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 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 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:

[[Page 7956]]

    (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 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 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 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 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 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 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 for this 
configuration does not exceed 110 percent of the V<INF>S 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.

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, 
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.
    (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 as defined in 
Sec. 23.73(a).

[[Page 7957]]

13. Stall Speed

    Instead of compliance with Sec. 23.49, the following apply:
    (a) V<INF>S 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 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 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 or more than 1.4 V<INF>S. 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 to V<INF>MO/M<INF>MO.)
    (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, 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; 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 to V<INF>MO/
M<INF>MO with the landing gear and flaps retracted, and from 1.4 
V<INF>S1 to VLE 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 
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/M<INF>FC, 
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, paragraph (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<SUP>.
    (b) Cruise. Static longitudinal stability must be shown in the 
cruise condition as follows:
    (1) With the landing gear retracted at high speed, the stick force 
curve must have a stable slope at all speeds within a range which is 
the greater of 15 percent of the trim speed plus the resulting free 
return speed range, or 50 knots plus the resulting free return speed 
range, above and below the trim speed (except that the speed range need 
not include speeds less than 1.4 V<INF>S1, nor speeds greater than 
V<INF>FC/M<INF>FC, 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/M<INF>MO; 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, 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 
+ 1.4 V<INF>S1)/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

[[Page 7958]]

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, nor speeds greater than 
V<INF>LE, 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; 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 and 1.8 V<INF>S1, 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 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 and 1.3 V<INF>S0 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 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 up to V<INF>FE, 
V<INF>LE, or V<INF>FC/M<INF>FC (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 or V<INF>LE, when appropriate) in the following airspeed 
ranges:
    (1) From 1.2 V<INF>S1 to V<INF>MO/M<INF>MO.
    (2) From V<INF>MO/M<INF>MO to V<INF>FC/M<INF>FC, 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 
(where V<INF>S1 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;
    (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

[[Page 7959]]

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/M<INF>DF. The maximum speeds shown must be 
used in establishing the operating limitations of the airplane in 
accordance with special condition 36.
    (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/M<INF>MO, 
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/M<INF>MO. 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/
M<INF>MO without the following:
    (i) Exceptional piloting strength or skill;
    (ii) Exceeding V<INF>D/M<INF>D, or V<INF>DF/M<INF>DF, 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/M<INF>DF with the airplane trimmed at V<INF>MO/M<INF>MO. 
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/M<INF>MO, the slope of the elevator 
control force versus speed curve need not be stable at speeds greater 
than V<INF>FC/M<INF>FC, but there must be a push force at all speeds up 
to V<INF>DF/M<INF>DF and there must be no sudden or excessive reduction 
of elevator control force as V<INF>DF/M<INF>DF is reached.
    (b) Maximum speed for stability characteristics. V<INF>FC/
M<INF>FC.. V<INF>FC/M<INF>FC 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/M<INF>MO and V<INF>DF/M<INF>DF except 
that, for altitudes where Mach number is the limiting factor, M<INF>FC 
need not exceed the Mach number at which effective speed warning 
occurs.

23. Flight Flutter Testing

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

24. Out-of-Trim Characteristics

    In the absence of specific requirements for out-of-trim 
characteristics, the Sino Swearingen Model SJ30-2 must comply with the 
following:
    (a) From an initial condition with the airplane trimmed at cruise 
speeds up to V<INF>MO/M<INF>MO, the airplane must have satisfactory 
maneuvering stability and controllability with the degree of out-of-
trim in both the airplane nose-up and nose-down directions, which 
results from the greater of the following:
    (1) A three-second movement of the longitudinal trim system at its 
normal rate for the particular flight condition with no aerodynamic 
load (or an equivalent degree of trim for airplanes that do not have a 
power-operated trim system), except as limited by stops in the trim 
system including those required by Sec. 23.655(b) for adjustable 
stabilizers; or
    (2) The maximum 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/M<INF>FC; and
    (2) At speeds between V<INF>FC/M<INF>FC and V<INF>DF/M<INF>DF, the 
direction of the primary longitudinal control force may not reverse.
    (c) Except as provided in paragraph (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

[[Page 7960]]

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/
M<INF>DF.
    (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/M<INF>DF 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/M<INF>DF 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 31 
(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

    In addition to the requirements of Sec. 23.607, the following apply 
to fasteners:
    (a) Each removable bolt, screw, nut, pin, or their removable 
fastener must incorporate two separate locking devices if the following 
apply:
    (1) Its loss could preclude continued flight and landing within the 
design limitations of the airplane using normal pilot skill and 
strength, or
    (2) Its loss could result in reduction in pitch, yaw, or roll 
control capability or response below that required by subpart B of this 
chapter and these special conditions.
    (b) The fasteners specified in paragraph (a) of this section and 
their locking devices may not be adversely affected by the 
environmental conditions associated with the particular installation.

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 operation 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 period, 
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 
period, 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

    In addition to the requirements of Sec. 23.841, the following 
apply--
    (a) The pressurization system--which includes, for this purpose, 
bleed air, air conditioning, and pressure control systems--must prevent 
the cabin altitude from exceeding the cabin altitude-time history shown 
in Figure 3 after each of the following:
    (1) Any probable malfunction or failure of the pressurization 
system. The existence of undetected, latent malfunctions or failures in 
conjunction with probable failures must be considered.
    (2) Any single failure in the pressurization system, combined with 
the occurrence of a leak produced by a complete loss of a door seal 
element, or a fuselage leak through an opening having an effective area 
2.0 times the effective area that produces the maximum permissible 
fuselage leak rate approved for normal operation, whichever produces a 
more severe leak.
    (b) The cabin altitude-time history may not exceed that shown in 
Figure 4 after each of the following:
    (1) The maximum pressure vessel opening resulting from an initially 
detectable crack propagating for a period encompassing four normal 
inspection intervals. Mid-panel cracks and cracks through skin-stringer 
and skin-frame combinations must be considered.
    (2) The pressure vessel opening or duct failure resulting from 
probable damage (failure effect) while under maximum operating cabin 
pressure differential due to a tire burst, engine rotor burst, loss of 
antennas or stall warning vanes, or any probable equipment failure 
(bleed air, pressure control, air conditioning, electrical sources(s), 
etc.) that affects pressurization.
    (3) Complete loss of thrust from all engines.
    (c) In showing compliance with paragraphs (a) and (b) of this 
special condition (Pressurization), it may be assumed that an emergency 
descent is made by approved emergency procedure. A seventeen-second 
flightcrew recognition and reaction time must be applied between cabin 
altitude

[[Page 7961]]

warning and the initiation of an emergency descent.

    Note: For the flight evaluation of the rapid descent, the test 
article must have the cabin volume representative of what is 
expected to be normal, such that Sino Swearingen must reduce the 
total cabin volume by that which would be occupied by the 
furnishings and total number of people.

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 to the maximum value 
of V<INF>2, 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.
    (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/M<INF>MO 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/M<INF>MO must be 
established so that it is not greater than the design cruising speed, 
V<INF>C, and so that it is sufficiently below V<INF>D/M<INF>D, or 
V<INF>DF/M<INF>DF, to make it highly improbable that the latter speeds 
will be inadvertently exceeded in operations. The speed margin between 
V<INF>MO/M<INF>MO and V<INF>D/M<INF>D, or V<INF>DF/M<INF>DF, 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 39, 40, and 41.
    (2) Other information that is necessary for safe operation because 
of design, operating, or handling characteristics.
    (3) Any limitation, procedure, or other information established as 
a condition of compliance with the applicable noise standards of Part 
36 of this chapter.
    (b) Approved Information. Each part of the manual listed in special 
conditions 39, 40, and 41, 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 which is readily 
accessible to the pilot.
    (f) Revisions and amendments. Each Airplane Flight Manual (AFM) 
must contain a means for recording the incorporation of revisions and 
amendments.

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/M<INF>MO, 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, and a statement that full 
application of rudder and aileron controls, as well as maneuvers that 
involve angles of attack

[[Page 7962]]

near the stall, should be confined to speeds below this value.
    (4) The maximum speed for flap extension, V<INF>FE, for the 
takeoff, approach, and landing positions.
    (5) The landing gear operating speed or speeds, V<INF>LO.
    (6) The landing gear extended speed, V<INF>LE if greater than 
V<INF>LO, 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 
Secs. 23.1549 through 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 
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 (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(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 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 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, paragraph 
(c), (d), and (e) that are related to the limitations and information 
required by paragraph (h) of special condition 39 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.


[[Page 7963]]


    Issued in Kansas City, Missouri on February 10, 1997.
Henry A. Armstrong,
Acting Manager, Small Airplane Directorate, Aircraft Certification 
Service.

BILLING CODE 4910-13-P
[GRAPHIC] [TIFF OMITTED] TP21FE97.011


[[Page 7964]]

[GRAPHIC] [TIFF OMITTED] TP21FE97.012



[FR Doc. 97-4353 Filed 2-20-97; 8:45 am]
BILLING CODE 4910-13-C