[Federal Register Volume 67, Number 126 (Monday, July 1, 2002)]
[Rules and Regulations]
[Pages 44018-44024]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 02-16386]


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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 25

[Docket No. NM213; Special Conditions No. 25-201-SC]


Special Conditions: Airbus, Model A340-500 and -600 Series 
Airplanes; Interaction of Systems and Structure; Electronic Flight 
Control System, Longitudinal Stability and Low Energy Awareness; and 
Use of High Incidence Protection and Alpha-Floor Systems

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Final special conditions.

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SUMMARY: These special conditions are issued for the Airbus Model A340-
500 and -600 series airplanes. These airplanes will have novel or 
unusual design features when compared to the state of technology 
envisioned in the airworthiness standards for transport category 
airplanes associated with the systems that affect the structural 
performance of the airplane; the electronic flight control system 
(EFCS); and the use of high incidence protection and alpha-floor 
systems. The applicable airworthiness regulations do not contain 
adequate or appropriate safety standards for these design features. 
These special conditions contain the additional safety standards that 
the Administrator considers necessary to establish a level of safety 
equivalent to that established by the existing airworthiness standards.

EFFECTIVE DATE: July 31, 2002.

FOR FURTHER INFORMATION CONTACT: Tim Backman, FAA, ANM-116, Transport 
Airplane Directorate, Aircraft Certification Service, 1601 Lind Avenue 
SW., Renton, Washington, 98055-4056; telephone (425) 227-2797; 
facsimile (425) 227-1149.

SUPPLEMENTARY INFORMATION:

Background

    On November 14, 1996, Airbus Industrie applied for an amendment to 
U.S. type certificate (TC) A43NM to include the new Models A340-500 and 
-600. These models are derivatives of the A340-300 airplane that is 
approved under the same TC.
    The Model A340-500 fuselage is a 6-frame stretch of the Model A340-
300 and is powered by 4 Rolls Royce Trent 553 engines; each rated at 
53,000 pounds of thrust. The airplane has interior seating arrangements 
for up to 375 passengers, with a maximum takeoff weight (MTOW) of 
820,000 pounds. The Model A340-500 is intended for long-range 
operations and has additional fuel capacity over that of the Model 
A340-600.
    The Model A340-600 fuselage is a 20-frame stretch of the Model 
A340-300 and is powered by 4 Rolls Royce Trent 556 engines; each rated 
at 56,000 pounds of thrust. The airplane has interior seating 
arrangements for up to 440 passengers, with a MTOW of 804,500 pounds.

Type Certification Basis

    Under the provisions of 14 CFR 21.101, Airbus must show that the 
Model A340-500 and -600 airplanes meet the applicable provisions of the 
regulations incorporated by reference in TC A43NM or the applicable 
regulations in effect on the date of application for the change to the 
type certificate. The regulations incorporated by reference in the type 
certificate are commonly referred to as the ``original type 
certification basis.'' The regulations incorporated by reference in TC 
A43NM are 14 CFR part 25, effective February 1, 1965, including 
Amendments 25-1 through 25-63, and Amendments 25-64, 25-65, 25-66, and 
25-77, with certain exceptions that are not relevant to these special 
conditions.
    In addition, if the regulations incorporated by reference do not 
provide adequate standards with respect to the change, the applicant 
must comply with certain regulations in effect on the date of 
application for the change. The FAA has determined that the Model A340-
500 and -600 airplanes must be shown to comply with Amendments 25-1 
through 25-91, and with certain FAA-allowed reversions for specific 
part 25 regulations to the part 25 amendment levels of the original 
type certification basis.
    Airbus has also chosen to comply with part 25 as amended by 
Amendments 25-92, -93, -94, -95, -97, -98, and -104. In addition, 
Airbus has elected to redefine the reference stall speed as the 1-g 
stall speed as proposed in Notice No. 95-17 (61 FR 1260, January 18, 
1996).
    If the Administrator finds that the applicable airworthiness 
regulations (i.e., part 25 as amended) do not contain adequate or 
appropriate safety standards for the Airbus Model A340-500 and ``600 
because of a novel or unusual design feature, special conditions are 
prescribed under the provisions of Sec. 21.16.
    In addition to the applicable airworthiness regulations and special 
conditions, the Airbus Model A340-500 and -600 must comply with the 
fuel vent and exhaust emission requirements of 14 CFR part 34 and the 
noise certification requirements of 14 CFR part 36, as amended on the 
date of type certification.
    Special conditions, as defined in 14 CFR 11.19, are issued in 
accordance with Sec. 11.38 and become part of the type certification 
basis in accordance with Sec. 21.101(b)(2).
    Special conditions are initially applicable to the model for which 
they are issued. Should the type certificate for that model be amended 
later to include any other model that incorporates the same novel or 
unusual design feature, or should any other model already included on 
the same type certificate be modified to incorporate the same novel or 
unusual design feature, the special conditions would also apply to the 
other model under the provisions of Sec. 21.101(a)(1).

Novel or Unusual Design Features

    The Airbus Model A340-500 and -600 airplanes will incorporate the 
following novel or unusual design features.

1. Interaction of Systems and Structure

    The Model A340-500 and -600 airplanes will have systems that affect 
the structural performance of the airplane, either directly or as a 
result of a failure or malfunction. These novel or unusual design 
features are systems that

[[Page 44019]]

can serve to alleviate loads in the airframe and, when in a failure 
state, can create loads in the airframe. The current regulations do not 
adequately account for the effects of these systems and their failures 
on structural performance. These special conditions provide the 
criteria to be used in assessing the effects of these systems on 
structures.

2. Electronic Flight Control System: Longitudinal Stability and Low 
Energy Awareness

    The EFCS of the Model A340-500 and -600, as with its predecessors, 
will result in the airplanes having neutral static longitudinal 
stability. This condition, when combined with the automatic trim 
feature of the EFCS, could result in insufficient feedback cues to the 
pilot of speed excursions below normal operating speeds. The 
longitudinal flight control laws provide neutral static stability 
within the normal flight envelope; therefore, the novel or unusual 
design features for these new airplane model designs will make them 
unable to show compliance with the static longitudinal stability 
requirements of Secs. 25.171, 25.173, and 25.175.
    The unique features of the Model A340-500 and -600 airplanes could 
cause an unsafe condition if the airspeed becomes too slow near the 
ground and results in the airplane stalling. The flightcrew would be 
unaware of the flight condition and would not be able to intervene and 
recover before stall. The French Direction Generale De L'Aviation 
Civile (DGAC) took action for this condition by introducing a special 
condition for predecessor airplanes with the same design features that 
required adequate awareness of the flightcrew to unsafe low speed 
conditions; there was no corresponding special condition developed by 
the FAA. The French special conditions allowed for awareness to be 
provided by an appropriate warning in the cockpit to allow for 
recovery. This special condition provides for an appropriate warning in 
the cockpit of the A340-500 and -600 airplanes to allow for recovery.
    Subsequent to certification of the predecessor Model A330 and A340 
airplanes and in establishing the certification requirements for the 
A340-500 and -600, the French DGAC decided to combine two special 
conditions from the A330 into a new special condition titled ``Static 
Longitudinal Stability and Low Energy Awareness.'' Since the FAA did 
not take action on the introduction of the low energy awareness 
requirement during the A330 and A340 certification, this special 
condition for the Model A340-500 and -600 airplane certification 
harmonizes to the French DGAC special condition for static longitudinal 
stability and low energy awareness. The purpose of the new low energy 
awareness special condition item 2(a)(2) is to provide awareness to the 
pilot of a low speed (or low energy state) of flight when the flight 
control laws provide neutral static longitudinal stability 
significantly below the normal operating speeds, and offer no cues to 
the pilot through the side stick controller. The special condition item 
2(a)(1) addresses the fact that the airplane has neutral stability and 
does not meet regulatory requirements for positive dynamic and static 
longitudinal stability (Secs. 25.171, 25.173, and 25.175, and 
25.181(a)).
3. High Incidence Protection and Alpha-floor Systems
    The Model A340-500 and -600 airplanes will have a novel or unusual 
feature to accommodate the unique features of the high incidence 
protection and the alpha-floor systems. The high incidence protection 
system replaces the stall warning system during normal operating 
conditions by prohibiting the airplane from stalling. The high 
incidence protection system limits the angle of attack at which the 
airplane can be flown during normal low speed operation, impacts the 
longitudinal airplane handling characteristics, and can not be over-
ridden by the crew. The existing regulations do not provide adequate 
criteria to address this system.
    The function of the alpha-floor system is to automatically increase 
the thrust on the operating engines under unusual circumstances where 
the airplane pitches to a predetermined high angle of attack or bank 
angle. The regulations do not provide adequate criteria to address this 
system.

Discussion of Comments

    Notice of proposed special conditions No. 25-02-05-SC for the 
Airbus Model A340-500 and -600 airplanes was published in the Federal 
Register on April 8, 2002 (67 FR 16656). No comments were received, and 
the special conditions are adopted as proposed.

Applicability

    As discussed above, these special conditions are applicable to the 
Model A340-500 and -600 airplanes. Should Airbus apply at a later date 
for a change to the type certificate to include another model 
incorporating the same novel or unusual design feature, these special 
conditions would apply to that model as well under the provisions of 
Sec. 21.101(a)(1).

Conclusion

    This action affects only certain novel or unusual design features 
on the Model A340-500 and -600 airplanes. It is not a rule of general 
applicability, and it affects only the applicant who applied to the FAA 
for approval of these features on the airplane.

List of Subjects in 14 CFR Part 25

    Aircraft, Aviation safety, Reporting and recordkeeping 
requirements.

    The authority citation for these special conditions is as follows:

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

The Special Conditions

    Accordingly, pursuant to the authority delegated by the 
Administrator, the following special conditions are issued as part of 
the type certification basis for Airbus Model A340-500 and -600 series 
airplanes.

1. Interaction of System and Structures

    The following special conditions are in lieu of compliance with the 
criteria of previously issued Special Conditions No. 25-ANM-69 (Docket 
No. NM-75), item 4, ``Interaction of Systems and Structure.''
    (a) General. For airplanes equipped with systems that affect 
structural performance, either directly or as a result of a failure or 
malfunction, the influence of these systems and their failure 
conditions must be taken into account when showing compliance with the 
requirements of subparts C and D of part 25. The following criteria 
must be used for showing compliance with these special conditions for 
airplanes equipped with flight control systems, autopilots, stability 
augmentation systems, load alleviation systems, flutter control 
systems, and fuel management systems. If these special conditions are 
used for other systems, it may be necessary to adapt the criteria to 
the specific system.
    (1) The criteria defined herein only address the direct structural 
consequences of the system responses and performances and cannot be 
considered in isolation but should be included in the overall safety 
evaluation of the airplane. These criteria may in some instances 
duplicate standards already established for this evaluation. These 
criteria are only applicable to structures whose failure could prevent 
continued safe flight and landing. Specific criteria that define 
acceptable limits on handling characteristics or

[[Page 44020]]

stability requirements when operating in the system degraded or 
inoperative modes are not provided in these special conditions.
    (2) Depending upon the specific characteristics of the airplane, 
additional studies that go beyond the criteria provided in these 
special conditions may be required in order to demonstrate the 
capability of the airplane to meet other realistic conditions; such as 
alternative gust or maneuver descriptions for an airplane equipped with 
a load alleviation system.
    (3) The following definitions are applicable to these special 
conditions.
    Structural performance: Capability of the airplane to meet the 
structural requirements of part 25.
    Flight limitations: Limitations that can be applied to the airplane 
flight conditions following an in-flight occurrence and that are 
included in the flight manual (e.g., speed limitations, avoidance of 
severe weather conditions, etc.).
    Operational limitations: Limitations, including flight limitations 
that can be applied to the airplane operating conditions before 
dispatch (e.g., fuel, payload, and Master Minimum Equipment List 
limitations).
    Probabilistic terms: The probabilistic terms (probable, improbable, 
extremely improbable) used in these special conditions are the same as 
those used in Sec. 25.1309.
    Failure condition: The term failure condition is the same as that 
used in Sec. 25.1309; however, these special conditions apply only to 
system failure conditions that affect the structural performance of the 
airplane (e.g., system failure conditions that induce loads, lower 
flutter margins, or change the response of the airplane to inputs such 
as gusts or pilot actions).
    (b) Effects of Systems on Structures. The following criteria will 
be used in determining the influence of a system and its failure 
conditions on the airplane structure.
    (1) System fully operative. With the system fully operative, the 
following apply:
    (i) Limit loads must be derived in all normal operating 
configurations of the system from all the limit conditions specified in 
subpart C, taking into account any special behavior of such a system or 
associated functions, or any effect on the structural performance of 
the airplane that may occur up to the limit loads. In particular, any 
significant nonlinearity (rate of displacement of control surface, 
thresholds or any other system nonlinearities) must be accounted for in 
a realistic or conservative way when deriving limit loads from limit 
conditions.
    (ii) The airplane must meet the strength requirements of part 25 
(static strength, residual strength), using the specified factors to 
derive ultimate loads from the limit loads defined above. The effect of 
nonlinearities must be investigated beyond limit conditions to ensure 
the behavior of the system presents no anomaly compared to the behavior 
below limit conditions. However, conditions beyond limit conditions 
need not be considered when it can be shown that the airplane has 
design features that will not allow it to exceed those limit 
conditions.
    (iii) The airplane must meet the aeroelastic stability requirements 
of Sec. 25.629.
    (2) System in the failure condition. For any system failure 
condition not shown to be extremely improbable, the following apply:
    (i) At the time of occurrence. Starting from 1-g level flight 
conditions, a realistic scenario, including pilot corrective actions, 
must be established to determine the loads occurring at the time of 
failure and immediately after failure.
    (A) For static strength substantiation, these loads multiplied by 
an appropriate factor of safety that is related to the probability of 
occurrence of the failure are ultimate loads to be considered for 
design. The factor of safety (FS) is defined in Figure 1.
[GRAPHIC] [TIFF OMITTED] TR01JY02.050

    (B) For residual strength substantiation, the airplane must be able 
to withstand two thirds of the ultimate loads defined in these special 
conditions item 1(b)(1)(ii).
    (C) Freedom from aeroelastic instability must be shown up to the 
speeds defined in Sec. 25.629(b)(2). For failure conditions that result 
in speed increases beyond Vc/Mc, freedom from aeroelastic instability 
must be shown to increased speeds, so that the margins intended by 
Sec. 25.629(b)(2) are maintained.
    (D) Failures of the system that result in forced structural 
vibrations (oscillatory failures) must not produce loads that could 
result in detrimental deformation of primary structure.
    (ii) For the continuation of the flight. For the airplane in the 
system failed state and considering any appropriate reconfiguration and 
flight limitations, the following apply:
    (A) The loads derived from the following conditions at speeds up to 
Vc, or the speed limitation prescribed for the remainder of the flight, 
must be determined:
    (1) The limit symmetrical maneuvering conditions specified in 
Sec. 25.331 and in Sec. 25.345.
    (2) The limit gust and turbulence conditions specified in 
Sec. 25.341 and in Sec. 25.345.

[[Page 44021]]

    (3) The limit rolling conditions specified in Sec. 25.349 and the 
limit unsymmetrical conditions specified in Sec. 25.367 and 
Sec. 25.427(b) and (c).
    (4) The limit yaw maneuvering conditions specified in Sec. 25.351.
    (5) The limit ground loading conditions specified in Sec. 25.473 
and Sec. 25.491.
    (B) For static strength substantiation, each part of the structure 
must be able to withstand the loads defined in special condition item 
1(b)(2)(ii)(A), multiplied by a factor of safety depending on the 
probability of being in this failure state. The factor of safety is 
defined in Figure 2.
[GRAPHIC] [TIFF OMITTED] TR01JY02.051

Qj = (Tj)(Pj) Where:

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

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

    (C) For residual strength substantiation, the airplane must be able 
to withstand two thirds of the ultimate loads defined in special 
condition item 1(b)(2)(ii)(B).
    (D) If the loads induced by the failure condition have a 
significant effect on fatigue or damage tolerance, then their effects 
must be taken into account.
    (E) Freedom from aeroelastic instability must be shown up to a 
speed determined from Figure 3. Flutter clearance speeds V\I\ and V\II\ 
may be based on the speed limitation specified for the remainder of the 
flight using the margins defined by Sec. 25.629(b).
[GRAPHIC] [TIFF OMITTED] TR01JY02.052

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

    Note to paragraph (E): If Pj is greater than 
10-\3\ per flight hour, then the flutter clearance speed 
must not be less than V\II\.

    (F) Freedom from aeroelastic instability must also be shown up to 
VI in Figure 3 above for any probable system failure 
condition combined with any damage required or selected for 
investigation by Sec. 25.571(b).
    (iii) Consideration of certain failure conditions may be required 
by other sections of part 25, regardless of calculated system 
reliability. Where analysis shows the probability of these failure 
conditions to be less than 10-9, criteria other than those 
specified in this paragraph may be used for structural substantiation 
to show continued safe flight and landing.
    (3) Warning considerations. For system failure detection and 
warning, the following apply:
    (i) The system must be checked for failure conditions, not 
extremely improbable, that degrade the structural capability below the 
level required by part 25 or significantly reduce the reliability of 
the remaining system. The flightcrew must be made aware of these 
failures before flight. Certain elements of the control system, such as 
mechanical and hydraulic components, may use special periodic 
inspections, and electronic components may use daily checks, in lieu of 
warning systems, to achieve the objective of this requirement. These 
certification maintenance requirements must be limited to components 
that are not readily detectable by normal warning

[[Page 44022]]

systems and where service history shows that inspections will provide 
an adequate level of safety.
    (ii) The existence of any failure condition, not shown to be 
extremely improbable, during flight that could significantly affect the 
structural capability of the airplane, and for which the associated 
reduction in airworthiness can be minimized by suitable flight 
limitations, must be signaled to the flightcrew. For example, failure 
conditions that result in a factor of safety between the airplane 
strength and the loads of subpart C below 1.25, or flutter margins 
below VII, must be signaled to the crew during flight.
    (4) Dispatch with known failure conditions. If the airplane is to 
be dispatched in a known system failure condition that affects 
structural performance, or affects the reliability of the remaining 
system to maintain structural performance, then the provisions of these 
special conditions must be met for the dispatched condition and for 
subsequent failures. Flight limitations and expected operational 
limitations may be taken into account in establishing Qj as the 
combined probability of being in the dispatched failure condition and 
the subsequent failure condition for the safety margins in Figures 2 
and 3. These limitations must be such that the probability of being in 
this combined failure state and then subsequently encountering limit 
load conditions is extremely improbable. No reduction in these safety 
margins is allowed if the subsequent system failure rate is greater 
than 10-3 per hour.

2. Electronic Flight Control System: Longitudinal Stability and Low 
Energy Awareness

    (a) The following special conditions are in lieu of compliance with 
the requirements of 14 CFR 25.171, 25.173, 25.175, and 25.181(a), and 
in lieu of compliance with the previously issued Special Conditions No. 
25-ANM-69 (Docket No. NM-75), item 11(b) ``Flight Characteristics--
Longitudinal Stability.''
    (1) The airplane must be shown to have suitable dynamic and static 
longitudinal stability in any condition normally encountered in 
service, including the effects of atmospheric disturbance.
    (2) The airplane must provide adequate awareness to the pilot of a 
low energy state when flight control laws provide neutral longitudinal 
stability significantly below the normal operating speeds.

3. High Incidence Protection and Alpha-Floor Systems

    (a) The following special conditions are in lieu of compliance with 
certain 14 CFR sections (listed below), and in lieu of compliance with 
previously issued Special Conditions No. 25-ANM-69 (Docket No. NM-75) 
item 12(b), ``Flight Envelope Protection, Angle-of-Attack Limiting.''
    (1) The following definitions are applicable to these special 
conditions.
    High Incidence Protection System. A system that operates directly 
and automatically on the airplane's flying controls to limit the 
maximum incidence that can be attained to a value below that at which 
an aerodynamic stall would occur.
    Alpha-floor System. A system that automatically increases thrust on 
the operating engines when incidence increases through a particular 
value.
    Alpha-limit. The maximum steady incidence at which the airplane 
stabilizes with the High Incidence Protection System operating and the 
longitudinal control held on its aft stop.
    Vmin. The minimum steady flight speed, for the airplane 
configuration under consideration and with the High Incidence 
Protection System operating, is the final stabilized Calibrated 
Airspeed obtained when the airplane is decelerated at an entry rate not 
exceeding 1 knot per second until the longitudinal pilot controller is 
on its stop.
    Vmin1g. Vmin corrected to 1g conditions. It 
is the minimum Calibrated Airspeed at which the airplane can develop a 
lift force normal to the flight path and equal to its weight when at an 
angle of attack not greater than that determined for Vmin.
    (2) Capability and Reliability of the High Incidence Protection 
System: In lieu of compliance with the requirements of previously 
issued Special Conditions No. 25-ANM-69, this special condition 
requires that acceptable capability and reliability of the High 
Incidence Protection System must be established by flight test, 
simulation, and analysis as appropriate. The capability and reliability 
required are as follows:
    (i) It shall not be possible during pilot induced maneuvers to 
encounter a stall and handling characteristics shall be acceptable, as 
required by special condition item 3(a)(5) of this special condition.
    (ii) The airplane shall be protected against stalling due to the 
effects of windshears and gusts at low speeds as required by special 
condition item 3(a)(6) of this special condition.
    (iii) The ability of the High Incidence Protection System to 
accommodate any reduction in stalling incidence resulting from residual 
ice must be verified.
    (iv) The reliability of the system and the effects of failures must 
be acceptable in accordance with Sec. 25.1309, and the associated 
policy.
    (3) Minimum Steady Flight Speed and Reference Stall Speed. In lieu 
of compliance with the requirements of Sec. 25.103 the following 
special conditions apply:
    (i) Vmin. The minimum steady flight speed, for the 
airplane configuration under consideration and with the High Incidence 
Protection System operating, is the final stabilized Calibrated 
Airspeed obtained when the airplane is decelerated at an entry rate not 
exceeding 1 knot per second until the longitudinal control is on its 
stop.
    (ii) The Minimum Steady Flight Speed, Vmin, must be 
determined with:
    (A) The High Incidence Protection System operating normally.
    (B) Idle thrust and Alpha-floor System inhibited.
    (C) All combinations of flap settings and landing gear positions.
    (D) The weight used when VSR is being used as a factor 
to determine compliance with a required performance standard.
    (E) The most unfavorable center of gravity allowable, and
    (F) The airplane trimmed for straight flight at a speed achievable 
by the automatic trim system.
    (iii) Vmin1g. Vmin corrected to 1g 
conditions. It is the minimum calibrated airspeed at which the airplane 
can develop a lift force normal to the flight path and equal to its 
weight when at an angle of attack not greater than that determined for 
Vmin. Vmin1g is defined as follows:
[GRAPHIC] [TIFF OMITTED] TR01JY02.053

where nZW = load factor normal to the flight path at 
Vmin
    (iv) The Reference Stall Speed, VSR, is a calibrated 
airspeed defined by the applicant. VSR may not be less than 
a 1-g stall speed. VSR is expressed as:
[GRAPHIC] [TIFF OMITTED] TR01JY02.054

where:
VCLMAX = Calibrated airspeed obtained when the load factor-
corrected lift coefficient
[GRAPHIC] [TIFF OMITTED] TR01JY02.055


is first a maximum during the maneuver prescribed in paragraph (v)(H) 
of this section.
nZW = Load factor normal to the flight path at 
VCLMAX
W = Airplane gross weight;
S = Aerodynamic reference wing area; and
q = Dynamic pressure.



[[Page 44023]]


    Note: Unless Angle of Attack (AOA) protection system (stall 
warning and stall identification) production tolerances are 
acceptably small, so as to produce insignificant changes in 
performance determinations, the flight test settings for stall 
warning and stall identification should be set at the low AOA 
tolerance limit; high AOA tolerance limits should be used for 
characteristics evaluations.


    (v) VSR must be determined with the following 
conditions:
    (A) Engines idling, or, if that resultant thrust causes an 
appreciable decrease in stall speed, not more than zero thrust at the 
stall speed.
    (B) The airplane in other respects (such as flaps and landing gear) 
in the condition existing in the test or performance standard in which 
VSR is being used.
    (C) The weight used when VSR is being used as a factor 
to determine compliance with a required performance standard.
    (D) The Center of gravity position that results in the highest 
value of reference stall speed.
    (E) The airplane trimmed for straight flight at a speed achievable 
by the automatic trim system, but not less than 1.13 VSR and 
not greater than 1.3 VSR.
    (F) The Alpha-floor system inhibited.
    (G) The High Incidence Protection System adjusted to a high enough 
incidence to allow full development of the 1g stall.
    (H) Starting from the stabilized trim condition, apply the 
longitudinal control to decelerate the airplane so that the speed 
reduction does not exceed one knot per second.
    (vi) The flight characteristics at the AOA for VCLMAX 
must be suitable in the traditional sense at FWD and AFT CG in straight 
and turning flight at IDLE power. Although for a normal production EFCS 
and steady full aft stick this AOA for VCLMAX cannot be 
achieved, the AOA can be obtained momentarily under dynamic 
circumstances and deliberately in a steady state sense with some EFCS 
failure conditions.
    (4) Stall Warning
    (i) Normal Operation. If the conditions of special conditions item 
3(a)(2) are satisfied, equivalent safety to the intent of Sec. 25.207, 
Stall Warning, shall be considered to have been met without provision 
of an additional, unique warning device.
    (ii) Failure Cases. Following failures of the High Incidence 
Protection System, not shown to be extremely improbable, such that the 
capability of the system no longer satisfies special conditions item 
3(a)(2)(i), (ii), and (iii), stall warning must be provided in 
accordance with Secs. 25.207(a), (b) and (f).
    (5) Handling Characteristics at High Incidence
    (i) High Incidence Handling Demonstrations. In lieu of compliance 
with the requirements of Sec. 25.201 the following apply:
    (A) Maneuvers to the limit of the longitudinal control, in the nose 
up direction, must be demonstrated in straight flight and in 30 degree 
banked turns with:
    (1) The high incidence protection system operating normally.
    (2) Initial power condition of:
    (i) Power off
    (ii) The power necessary to maintain level flight at 1.5 
VSR1, where VSR1 is the stall speed with the 
flaps in the approach position, the landing gear retracted, and the 
maximum landing weight. The flap position to be used to determine this 
power setting is that position in which the stall speed, 
VSR1, does not exceed 110 percent of the stall speed, 
VSR0, with the flaps in the most extended landing position.
    (3) Alpha-floor system operating normally unless more severe 
conditions are achieved with alpha-floor inhibited.
    (4) Flaps, landing gear and deceleration devices in any likely 
combination of positions.
    (5) Representative weights within the range for which certification 
is requested, and
    (6) The airplane trimmed for straight flight at a speed achievable 
by the automatic trim system.
    (B) The following procedures must be used to show compliance with 
the requirements of special condition item 3(a)(5)(ii).
    (1) Starting at a speed sufficiently above the minimum steady 
flight speed to ensure that a steady rate of speed reduction can be 
established, apply the longitudinal control so that the speed reduction 
does not exceed one knot per second until the control reaches the stop.
    (2) The longitudinal control must be maintained at the stop until 
the airplane has reached a stabilized flight condition and must then be 
recovered by normal recovery techniques.
    (3) The requirements for turning flight maneuver demonstrations 
must also be met with accelerated rates of entry to the incidence 
limit, up to the maximum rate achievable.
    (ii) Characteristics in High Incidence Maneuvers. In lieu of 
compliance with the requirements of Sec. 25.203, the following apply:
    (A) Throughout maneuvers with a rate of deceleration of not more 
than 1 knot per second, both in straight flight and in 30 degree banked 
turns, the airplane's characteristics shall be as follows:
    (1) There shall not be any abnormal airplane nose-up pitching.
    (2) There shall not be any uncommanded nose-down pitching, which 
would be indicative of stall. However, reasonable attitude changes 
associated with stabilizing the incidence at alpha limit as the 
longitudinal control reaches the stop would be acceptable. Any 
reduction of pitch attitude associated with stabilizing the incidence 
at the alpha limit should be achieved smoothly and at a low pitch rate, 
such that it is not likely to be mistaken for natural stall 
identification.
    (3) There shall not be any uncommanded lateral or directional 
motion, and the pilot must retain good lateral and directional control, 
by conventional use of the cockpit controllers, throughout the 
maneuver.
    (4) The airplane must not exhibit severe buffeting of a magnitude 
and severity that would act as a deterrent to completing the maneuver.
    (B) In maneuvers with increased rates of deceleration, some 
degradation of characteristics is acceptable, associated with a 
transient excursion beyond the stabilized Alpha-limit. However, the 
airplane must not exhibit dangerous characteristics or characteristics 
that would deter the pilot from holding the longitudinal controller on 
the stop for a period of time appropriate to the maneuvers.
    (C) It must always be possible to reduce incidence by conventional 
use of the controller.
    (D) The rate at which the airplane can be maneuvered from trim 
speeds associated with scheduled operating speeds such as V2 
and Vref up to Alpha-limit shall not be unduly damped or 
significantly slower than can be achieved on conventionally controlled 
transport airplanes.
    (6) Atmospheric Disturbances.
    Operation of the High Incidence Protection System and the Alpha-
floor System must not adversely affect aircraft control during expected 
levels of atmospheric disturbances, nor impede the application of 
recovery procedures in case of windshear. Simulator tests and analysis 
may be used to evaluate such conditions, but must be validated by 
limited flight testing to confirm handling qualities at critical 
loading conditions.
    (7) Alpha Floor.
    The Alpha-floor setting must be such that the aircraft can be flown 
at normal landing operational speed and maneuvered up to bank angles 
consistent with the flight phase (including the maneuver capabilities 
specified in Sec. 25.143(g)) of the 1-g stall Equivalent Safety Finding 
without

[[Page 44024]]

triggering Alpha-floor. In addition, there must be no Alpha-floor 
triggering unless appropriate when the airplane is flown in usual 
operational maneuvers and in turbulence.
    (8) In lieu of compliance with the requirements of Sec. 25.145, the 
following apply:
    (i) It must be possible, at any point between the trim speed 
prescribed in special condition item 3(a)(ii)(F), and Vmin, 
to pitch the nose downward so that the acceleration to this selected 
trim speed is prompt with:
    (ii) The airplane trimmed at the trim speed prescribed in special 
condition item 3(a)(ii)(F);
    (A) The landing gear extended;
    (B) The wing flaps retracted and extended; and
    (C) Power off and at maximum continuous power on the engines.
    (9) In lieu of compliance with the requirements of 
Sec. 25.145(b)(6), the following apply:
    With power off, flaps extended and the airplane trimmed at 1.3 
VSR1, obtain and maintain airspeeds between Vmin 
and either 1.6VSR1 or VFE, whichever is lower.
    (10) In lieu of compliance with the requirements of 
Sec. 25.1323(c), the following apply:
    (i) VMO to Vmin with the flaps retracted; and
    (ii) Vmin to VFE with flaps in the landing 
position.

    Issued in Renton, Washington, on June 17, 2002.
Kalene C. Yanamura,
Acting Manager, Transport Airplane Directorate, Aircraft Certification 
Service.
[FR Doc. 02-16386 Filed 6-28-02; 8:45 am]
BILLING CODE 4910-13-P