[Federal Register: April 11, 2006 (Volume 71, Number 69)]
[Rules and Regulations]
[Page 18169-18183]
From the Federal Register Online via GPO Access [wais.access.gpo.gov]
[DOCID:fr11ap06-4]
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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 25
[Docket No. NM305; Special Conditions No. 25-316-SC]
Special Conditions: Airbus Model A380-800 Airplane; Dynamic
Braking, Interaction of Systems And Structures, Limit Pilot Forces,
Side Stick Controllers, Dive Speed Definition, Electronic Flight
Control System-Lateral-Directional Stability, Longitudinal Stability,
And Low Energy Awareness, Electronic Flight Control System-Control
Surface Awareness, Electronic Flight Control System-Flight
Characteristics Compliance Via the Handling Qualities Rating Method,
Flight Envelope Protection-General Limiting Requirements, Flight
Envelope Protection-Normal Load Factor (G) Limiting, Flight Envelope
Protection-High Speed Limiting, Flight Envelope Protection-Pitch And
Roll Limiting, Flight Envelope Protection-High Incidence Protection and
Alpha-Floor Systems, High Intensity Radiated Fields (HIRF) Protection,
and Operation Without Normal Electrical Power
AGENCY: Federal Aviation Administration (FAA), DOT.
ACTION: Final Special Conditions.
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SUMMARY: These Special Conditions are issued for the Airbus A380-800
airplane. This airplane will have novel or unusual design features when
compared to the state of technology envisioned in the airworthiness
standards for transport category airplanes. These design features
include side stick controllers, a body landing gear in addition to
conventional wing and nose landing gears, electronic flight control
systems, and flight envelope protection. These Special Conditions also
pertain to the effects of such novel or unusual design features, such
as their effects on the structural performance of the airplane.
Finally, the Special Conditions pertain to the effects of certain
conditions on these novel or unusual design features, such as the
effects of high intensity radiated fields (HIRF) or of operation
without normal electrical power. Additional Special Conditions will be
issued for other novel or unusual design features of the Airbus A380-
800 airplanes. A list is provided in the section of this document
entitled ``Discussion of Novel or Unusual Design Features.''
EFFECTIVE DATE: March 30, 2006.
FOR FURTHER INFORMATION CONTACT: Holly Thorson, FAA, International
Branch, ANM-116, Transport Airplane Directorate, Aircraft Certification
Service, 1601 Lind Avenue, SW., Renton, Washington 98055-4056;
telephone (425) 227-1357; facsimile (425) 227-1149.
SUPPLEMENTARY INFORMATION
Background
Airbus applied for FAA certification/validation of the
provisionally-designated Model A3XX-100 in its letter AI/L 810.0223/98,
dated August 12, 1998, to the FAA. Application for certification by the
Joint Aviation Authorities (JAA) of Europe had been made on January 16,
1998, reference AI/L 810.0019/98. In its letter to the FAA, Airbus
requested an extension to the 5-year period for type certification in
accordance with 14 CFR 21.17(c).
The request was for an extension to a 7-year period, using the date
of the initial application letter to the JAA as the reference date. The
reason given by Airbus for the request for extension is related to the
technical challenges, complexity, and the number of new and novel
features on the airplane. On November 12, 1998, the Manager, Aircraft
Engineering Division, AIR-100, granted Airbus' request for the 7-year
period based on the date of application to the JAA.
In its letter AI/LE-A 828.0040/99 Issue 3, dated July 20, 2001,
Airbus stated that its target date for type certification of the Model
A380-800 had been moved from May 2005, to January 2006, to match the
delivery date of the first production airplane. In a subsequent letter
(AI/L 810.0223/98 issue 3, dated January 27, 2006), Airbus stated that
its target date for type certification is October 2, 2006. In
accordance with 14 CFR 21.17(d)(2), Airbus chose a new application date
of December 20, 1999, and requested that the 7-year certification
period which had already been approved be continued. The FAA has
reviewed the part 25 certification basis for the Model A380-800
airplane, and no changes are required based on the new application
date.
The Model A380-800 airplane will be an all-new, four-engine jet
transport airplane with a full double-deck, two-aisle cabin. The
maximum takeoff weight will be 1.235 million pounds with a typical
three-class layout of 555 passengers.
Type Certification Basis
Under the provisions of 14 CFR 21.17, Airbus must show that the
Model A380-800 airplane meets the applicable provisions of 14 CFR part
25, as amended by Amendments 25-1 through 25-98. If the Administrator
finds that the applicable airworthiness regulations do not contain
adequate or appropriate safety standards for the Airbus A380-800
airplane because of novel or unusual design features, Special
Conditions are prescribed under the provisions of 14 CFR 21.16.
In addition to the applicable airworthiness regulations and Special
Conditions, the Airbus Model A380-800 airplane must comply with the
fuel vent and exhaust emission requirements of 14 CFR part 34 and the
noise certification requirements of 14 CFR part 36. In addition, the
FAA must issue a finding of regulatory adequacy pursuant to section 611
of Public Law 93-574, the ``Noise Control Act of 1972.''
Special Conditions, as defined in 14 CFR 11.19, are issued in
accordance with 14 CFR 11.38 and become part of the type certification
basis in accordance with 14 CFR 21.17(a)(2).
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, the Special Conditions would also apply to the
other model under the provisions of 14 CFR 21.101.
Discussion of Novel or Unusual Design Features
The Airbus A380-800 airplane will incorporate a number of novel or
unusual design features. Because of rapid improvements in airplane
technology, the applicable airworthiness regulations do not contain
adequate or appropriate safety standards for these design features.
These Special Conditions for Airbus Model A380 contain the
additional safety standards that the Administrator considers necessary
to establish a level of safety equivalent to that established by the
existing airworthiness standards.
[[Page 18170]]
These Special Conditions are identical or nearly identical to those
previously required for type certification of the basic Model A340
airplane or earlier models. One exception is the Special Conditions
pertaining to Interaction of Systems and Structures. It was not
required for the basic Model A340 but was required for type
certification of the larger, heavier Model A340-500 and--600 airplanes.
In general, the Special Conditions were derived initially from
standardized requirements developed by the Aviation Rulemaking Advisory
Committee (ARAC), comprised of representatives of the FAA, Europe's
Joint Aviation Authorities (now replaced by the European Aviation
Safety Agency), and industry. In some cases, a draft Notice of Proposed
Rulemaking has been prepared but no final rule has yet been
promulgated.
Additional Special Conditions will be issued for other novel or
unusual design features of the Airbus Model A380-800 airplane. Those
Special Conditions pertain to the following topics:
Fire protection,
Evacuation, including availability of stairs in an
emergency,
Emergency exit arrangement--outside viewing,
Escape system inflation systems,
Escape systems installed in non-pressurized compartments,
Ground turning loads,
Crashworthiness,
Flotation and ditching,
Discrete gust requirements,
Transient engine failure loads,
Airplane jacking loads,
Landing gear pivoting loads,
Design roll maneuvers,
Extendable length escape systems,
Reinforced flightdeck bulkhead, and
Lithium ion battery installations.
1. Dynamic Braking
The A380 landing gear system will include body gear in addition to
the conventional wing and nose gear. This landing gear configuration
may result in more complex dynamic characteristics than those found in
conventional landing gear configurations. Section 25.493(d) by itself
does not contain an adequate standard for assessing the braking loads
for the A380 landing gear configuration.
Due to the potential complexities of the A380 landing gear system,
in addition to meeting the requirements of Sec. 25.493(d), a rational
analysis of the braked roll conditions is necessary. Airbus Model A340-
500 and -600 also have a body-mounted main landing gear in addition to
the wing and nose gears. Therefore, Special Conditions similar to those
required for that model are appropriate for the model A380-800.
2. Interaction of Systems and Structures
The A380 is equipped with systems which affect the airplane's
structural performance either directly or as a result of failure or
malfunction. The effects of these systems on structural performance
must be considered in the certification analysis. This analysis must
include consideration of normal operation and of failure conditions
with required structural strength levels related to the probability of
occurrence.
Previously, Special Conditions have been specified to require
consideration of the effects of systems on structures. The Special
Conditions for the Model A380 are nearly identical to those issued for
the Model A340-500 and -600 series airplanes.
3. Limit Pilot Forces
Like some other Airbus models, the Model A380 airplane is equipped
with a side stick controller instead of a conventional control stick.
This kind of controller is designed to be operated using only one hand.
The requirement of Sec. 25.397(c), which defines limit pilot forces
and torques for conventional wheel or stick controls, is not
appropriate for a side stick controller. Therefore, Special Conditions
are necessary to specify the appropriate loading conditions for this
kind of controller.
Special Conditions for side stick controllers have already been
developed for the Airbus model A320 and A340 airplanes, both of which
also have a side stick controller instead of a conventional control
stick. The same Special Conditions are appropriate for the model A380
airplane.
4. Side Stick Controllers
The A380--like its predecessors, the A320, A330, and A340--will use
side stick controllers for pitch and roll control. Regulatory
requirements for conventional wheel and column controllers, such as
requirements pertaining to pilot strength and controllability, are not
directly applicable to side stick controllers. In addition, pilot
control authority may be uncertain, because the side sticks are not
mechanically interconnected as with conventional wheel and column
controls.
In previous Airbus airplane certification programs, Special
Conditions pertaining to side stick controllers were addressed in three
separate issue papers, entitled ``Pilot Strength,'' ``Pilot Coupling,''
and ``Pilot Control.'' The resulting separate Special Conditions are
combined in these Special Conditions under the title of ``Side Stick
Controllers.'' In order to harmonize with the JAA, the following has
been added to Special Conditions 4.c. Side Stick Controllers:
Pitch and roll control force and displacement sensitivity must be
compatible, so that normal inputs on one control axis will not cause
significant unintentional inputs on the other.
5. Dive Speed Definition
Airbus proposes to reduce the speed spread between VC
and VD required by Sec. 25.335(b), based on the
incorporation of a high speed protection system in the A380 flight
control laws. The A380--like the A320, A330, and A340--is equipped with
a high speed protection system which limits nose down pilot authority
at speeds above VC/MC and prevents the airplane
from actually performing the maneuver required under Sec.
25.335(b)(1).
Section 25.335(b)(1) is an analytical envelope condition which was
originally adopted in Part 4b of the Civil Air Regulations to provide
an acceptable speed margin between design cruise speed and design dive
speed. Freedom from flutter and airframe design loads is affected by
the design dive speed. While the initial condition for the upset
specified in the rule is 1g level flight, protection is afforded for
other inadvertent overspeed conditions as well. Section 25.335(b)(1) is
intended as a conservative enveloping condition for all potential
overspeed conditions, including non-symmetric ones. To establish that
all potential overspeed conditions are enveloped, the applicant must
demonstrate either of the following:
Any reduced speed margin--based on the high speed
protection system in the A380--will not be exceeded in inadvertent or
gust induced upsets, resulting in initiation of the dive from non-
symmetric attitudes; or
The airplane is protected by the flight control laws from
getting into non-symmetric upset conditions.
In addition, the high speed protection system in the A380 must have
a high level of reliability.
6. Electronic Flight Control System: Lateral-Directional Stability,
Longitudinal Stability, and Low Energy Awareness
In lieu of compliance with the regulations pertaining to lateral-
directional and longitudinal stability, these Special Conditions ensure
that the model A380 will have suitable airplane
[[Page 18171]]
handling qualities throughout the normal flight envelope (reference
paragraphs 6.a. and 6.b.).
The unique features of the A380 flight control system and side-
stick controllers, when compared with conventional airplanes with wheel
and column controllers, do not provide conventional awareness to the
flight crew of a change in speed or a change in the direction of flight
(reference paragraph 6.c.). These Special Conditions requires that
adequate awareness be provided to the pilot of a low energy state (low
speed, low thrust, and low altitude) below normal operating speeds.
a. Lateral-directional Static Stability: The model A380 airplane
has a flight control design feature within the normal operational
envelope in which side stick deflection in the roll axis commands roll
rate. As a result, the stick force in the roll axis will be zero
(neutral stability) during the straight, steady sideslip flight
maneuver of Sec. 25.177(c) and will not be ``substantially
proportional to the angle of sideslip,'' as required by the regulation.
The electronic flight control system (EFCS) on the A380 as on its
predecessors--the A320, A330 and A340--contains fly-by-wire control
laws that result in neutral lateral-directional static stability.
Therefore, the conventional requirements of the regulations are not
met.
With conventional control system requirements, positive static
directional stability is defined as the tendency to recover from a skid
with the rudder free. Positive static lateral stability is defined as
the tendency to raise the low wing in a sideslip with the aileron
controls free. The regulations are intended to accomplish the
following:
Provide additional cues of inadvertent sideslips and skids
through control force changes.
Ensure that short periods of unattended operation do not
result in any significant changes in yaw or bank angle.
Provide predictable roll and yaw response.
Provide acceptable level of pilot attention (i.e.,
workload) to attain and maintain a coordinated turn.
b. Longitudinal Static Stability: The longitudinal flight control
laws for the A380 provide neutral static stability within the normal
operational envelope. Therefore, the airplane design does not comply
with the static longitudinal stability requirements of Sec. Sec.
25.171, 25.173, and 25.175.
Static longitudinal stability on conventional airplanes with
mechanical links to the pitch control surface means that a pull force
on the controller will result in a reduction in speed relative to the
trim speed, and a push force will result in higher than trim speed.
Longitudinal stability is equired by the regulations for the following
reasons:
Speed change cues are provided to the pilot through
increased and decreased forces on the controller.
Short periods of unattended control of the airplane do not
result in significant changes in attitude, airspeed, or load factor.
A predictable pitch response is provided to the pilot.
An acceptable level of pilot attention (i.e., workload) to
attain and maintain trim speed and altitude is provided to the pilot.
Longitudinal stability provides gust stability.
The pitch control movement of the side stick is a normal load
factor or ``g'' command which results in an initial movement of the
elevator surface to attain the commanded load factor. That movement is
followed by integrated movement of the stabilizer and elevator to
automatically trim the airplane to a neutral (1g) stick-free stability.
The flight path commanded by the initial side stick input will remain
stick-free until the pilot gives another command. This control function
is applied during ``normal'' control law within the speed range from
V[alpha]prot (the speed at the angle of attack protection
limit) to VMO/MMO. Once outside this speed range,
the control laws introduce the conventional longitudinal static
stability as described above.
As a result of neutral static stability, the A380 does not meet the
requirements of part 25 for static longitudinal stability.
c. Low Energy Awareness: Static longitudinal stability provides an
awareness to the flight crew of a low energy state (low speed and
thrust at low altitude). Past experience on airplanes fitted with a
flight control system which provides neutral longitudinal stability
shows there are insufficient feedback cues to the pilot of excursion
below normal operational speeds. The maximum angle of attack protection
system limits the airplane angle of attack and prevents stall during
normal operating speeds, but this system is not sufficient to prevent
stall at low speed excursions below normal operational speeds. Until
intervention, there are no stability cues, because the airplane remains
trimmed. Additionally, feedback from the pitching moment due to thrust
variation is reduced by the flight control laws. Recovery from a low
speed excursion may become hazardous when the low speed is associated
with low altitude and the engines are operating at low thrust or with
other performance limiting conditions.
7. Electronic Flight Control System: Control Surface Awareness
With a response-command type of flight control system and no direct
coupling from cockpit controller to control surface, such as on the
A380, the pilot is not aware of the actual surface deflection position
during flight maneuvers. Some unusual flight conditions, arising from
atmospheric conditions or airplane or engine failures or both, may
result in full or nearly full surface deflection. Unless the flight
crew is made aware of excessive deflection or impending control surface
deflection limiting, piloted or auto-flight system control of the
airplane might be inadvertently continued in a way which would cause
loss of control or other unsafe handling or performance
characteristics.
These Special Conditions requires that suitable annunciation be
provided to the flight crew when a flight condition exists in which
nearly full control surface deflection occurs. Suitability of such a
display must take into account that some pilot-demanded maneuvers
(e.g., rapid roll) are necessarily associated with intended full or
nearly full control surface deflection. Therefore, simple alerting
systems which would function in both intended or unexpected control-
limiting situations must be properly balanced between needed crew
awareness and not getting nuisance warnings.
8. Electronic Flight Control System: Flight Characteristics Compliance
Via the Handling Qualities Rating Method (HQRM)
The Model A380 airplane will have an Electronic Flight Control
System (EFCS). This system provides an electronic interface between the
pilot's flight controls and the flight control surfaces (for both
normal and failure states). The system also generates the actual
surface commands that provide for stability augmentation and control
about all three airplane axes. Because EFCS technology has outpaced
existing regulations--written essentially for unaugmented airplanes
with provision for limited ON/OFF augmentation--suitable Special
Conditions and a method of compliance are required to aid in the
certification of flight characteristics.
These Special Conditions and the method of compliance presented in
Appendix 7 of the Flight Test Guide, AC 25-7A, provide a means by which
one may evaluate flight characteristics--as,
[[Page 18172]]
for example, ``satisfactory,'' ``adequate,'' or ``controllable''--to
determine compliance with the regulations. The HQRM in Appendix 7 was
developed for airplanes with control systems having similar functions
and is employed to aid in the evaluation of the following:
All EFCS/airplane failure states not shown to be extremely
improbable and where the envelope (task) and atmospheric disturbance
probabilities are each 1.
All combinations of failures, atmospheric disturbance
level, and flight envelope not shown to be extremely improbable.
The HQRM provides a systematic approach to the assessment of
handling qualities. It is not intended to dictate program size or need
for a fixed number of pilots to achieve multiple opinions. The airplane
design itself and success in defining critical failure combinations
from the many reviewed in Systems Safety Assessments would dictate the
scope of any HQRM application.
Handling qualities terms, principles, and relationships familiar to
the aviation community have been used to formulate the HQRM. For
example, we have established that the well-known COOPER-HARPER rating
scale and the proposed FAA three-part rating system are similar. This
approach is derived in part from the contract work on the flying
qualities of highly augmented/ relaxed static stability airplanes, in
relation to regulatory and flight test guide requirements. The work is
reported in DOT/FAA/CT-82/130, Flying Qualities of Relaxed Static
Stability Aircraft, Volumes I and II.
9. Flight Envelope Protection: General Limiting Requirements
These Special Conditions and the following ones--pertaining to
flight envelope protection--present general limiting requirements for
all the unique flight envelope protection features of the basic A380
Electronic Flight Control System (EFCS) design. Current regulations do
not address these types of protection features. The general limiting
requirements are necessary to ensure a smooth transition from normal
flight to the protection mode and adequate maneuver capability. The
general limiting requirements also ensure that the structural limits of
the airplane are not exceeded. Furthermore, failure of the protection
feature must not create hazardous flight conditions. Envelope
protection parameters include angle of attack, normal load factor, bank
angle, pitch angle, and speed. To accomplish these envelope
protections, one or more significant changes occur in the EFCS control
laws as the normal flight envelope limit is approached or exceeded.
Each specific type of envelope protection is addressed individually
in the Special Conditions which follow.
10. Flight Envelope Protection: Normal Load Factor (G) Limiting
The A380 flight control system design incorporates normal load
factor limiting on a full time basis that will prevent the pilot from
inadvertently or intentionally exceeding the positive or negative
airplane limit load factor. This limiting feature is active in all
normal and alternate flight control modes and cannot be overridden by
the pilot. There is no requirement in the regulations for this limiting
feature.
Except for the Airbus airplanes with fly-by-wire flight controls,
the normal load factor limit is unique in that traditional airplanes
with conventional flight control systems (mechanical linkages) are
limited in the pitch axis only by the elevator surface area and
deflection limit. The elevator control power is normally derived for
adequate controllability and maneuverability at the most critical
longitudinal pitching moment. The result is that traditional airplanes
have a significant portion of the flight envelope in which
maneuverability in excess of limit structural design values is
possible.
Part 25 does not require a demonstration of maneuver control or
handling qualities beyond the design limit structural loads.
Nevertheless, some pilots have become accustomed to the availability of
this excess maneuver capacity in case of extreme emergency, such as
upset recoveries or collision avoidance. Airbus is aware of the concern
and has published the results of its research which indicate the
following:
Pilots rarely, if ever, use the excess maneuvering
capacity in collision avoidance maneuvers, and
Other features of its flight control system would have
prevented most, if not all, of the upset cases on record where pilots
did exceed limit loads during recovery.
Because Airbus has chosen to include this optional design feature
for which part 25 does not contain adequate or appropriate safety
standards, Special Conditions pertaining to this feature are included.
These Special Conditions establish minimum load factor requirements to
ensure adequate maneuver capability during normal flight. Other
limiting features of the normal load factor limiting function, as
discussed above, that would affect the upper load limits are not
addressed in these Special Conditions. The phrase ``in the absence of
other limiting factors'' has been added relative to past similar
Special Condition to clarify that while the main focus is on the lower
load factor limits, there are other limiting factors that must be
considered in the load limiting function.
11. Flight Envelope Protection: High Speed Limiting
The longitudinal control law design of the A380 incorporates a high
speed limiting protection system in the normal flight mode. This system
prevents the pilot from inadvertently or intentionally exceeding the
airplane maximum design speeds, VD MD. Part 25
does not address such a system that would limit or modify flying
qualities in the high speed region.
The main features of the high speed limiting function are as
follows:
It protects the airplane against high speed/high mach
number flight conditions beyond VMO/MMO.
It does not interfere with flight at VMO/
MMO, even in turbulent air.
It still provides load factor limitation through the
``pitch limiting'' function described below.
It restores positive static stability beyond
VMO/MMO.
This Special Condition establishes requirements to ensure that
operation of the high speed limiter does not impede normal attainment
of speeds up to the overspeed warning.
12. Flight Envelope Protection: Pitch and Roll Limiting
Currently, part 25 does not specifically address flight
characteristics associated with fixed attitude limits. Airbus proposes
to implement pitch and roll attitude limiting functions on the A380 via
the Electronic Flight Control System (EFCS) normal modes. These normal
modes will prevent airplane pitch attitudes greater than +30 degrees
and less than -15 degrees and roll angles greater than plus or minus 67
degrees. In addition, positive spiral stability is introduced for roll
angles greater than 33 degrees at speeds below VMO/
MMO. At speeds greater than VMO/MMO,
the maximum aileron control force with positive spiral stability
results in a maximum bank angle of 45 degrees.
These Special Conditions establish requirements to ensure that
pitch limiting functions do not impede normal maneuvering and that
pitch and roll limiting functions do not restrict or prevent attaining
certain roll angles necessary for emergency maneuvering.
Special Conditions to supplement Sec. 25.143 concerning pitch and
roll limits
[[Page 18173]]
were developed for the A320, A330 and A340 in which performance of the
limiting functions was monitored throughout the flight test program.
The FAA expects similar monitoring to take place during the A380 flight
test program to substantiate the pitch and roll attitude limiting
functions and the appropriateness of the chosen limits.
13. Flight Envelope Protection: High Incidence Protection and Alpha-
floor Systems
The A380 is equipped with a high incidence protection system that
limits the angle of attack at which the airplane can be flown during
normal low speed operation and that cannot be overridden by the flight
crew. The application of this limitation on the angle of attack affects
the longitudinal handling characteristics of the airplane, so that
there is no need for the stall warning system during normal operation.
In addition, the alpha-floor function automatically advances the
throttles on the operating engines whenever the airplane angle of
attack reaches a predetermined high value. This function is intended to
provide increased climb capability. This Special Conditions thus
addresses the unique features of the low speed high incidence
protection and the alpha-floor systems on the A380.
The high incidence protection system prevents the airplane from
stalling, which means that the stall warning system is not needed
during normal flight conditions. If there is a failure of the high
incidence protection system that is not shown to be extremely
improbable, the flight characteristics at the angle of attack for
CLMAX must be suitable in the traditional sense, and stall
warning must be provided in a conventional manner.
14. High Intensity Radiated Fields (HIRF) Protection
The Airbus Model A380-800 will utilize electrical and electronic
systems which perform critical functions. These systems may be
vulnerable to high-intensity radiated fields (HIRF) external to the
airplane. There is no specific regulation that addresses requirements
for protection of electrical and electronic systems from HIRF. With the
trend toward increased power levels from ground-based transmitters and
the advent of space and satellite communications, coupled with
electronic command and control of the airplane, the immunity of
critical avionics/electronics and electrical systems to HIRF must be
established.
To ensure that a level of safety is achieved that is equivalent to
that intended by the regulations incorporated by reference, Special
Conditions are needed for the Airbus Model A380 airplane. These Special
Conditions require that avionics/electronics and electrical systems
that perform critical functions be designed and installed to preclude
component damage and interruption.
It is not possible to precisely define the HIRF to which the
airplane will be exposed in service. There is also uncertainty
concerning the effectiveness of airframe shielding for HIRF.
Furthermore, coupling of electromagnetic energy to cockpit-installed
equipment through the cockpit window apertures is undefined. Based on
surveys and analysis of existing HIRF emitters, adequate protection
from HIRF exists when there is compliance with either paragraph a. or
b. below:
a. A minimum threat of 100 volts rms (root-mean-square) per meter
electric field strength from 10 KHz to 18 GHz.
(1) The threat must be applied to the system elements and their
associated wiring harnesses without the benefit of airframe shielding.
(2) Demonstration of this level of protection is established
through system tests and analysis.
b. A threat external to the airframe of the field strengths
indicated in the table below for the frequency ranges indicated. Both
peak and average field strength components from the table below are to
be demonstrated.
------------------------------------------------------------------------
Field strength
(volts per meter)
Frequency -------------------
Peak Average
------------------------------------------------------------------------
10 kHz-100 kHz...................................... 50 50
100 kHz-500 kHz..................................... 50 50
500 kHz-2 MHz....................................... 50 50
2 MHz-30 MHz........................................ 100 100
30 MHz-70 MHz....................................... 50 50
70 MHz-100 MHz...................................... 50 50
100 MHz-200 MHz..................................... 100 100
200 MHz-400 MHz..................................... 100 100
400 MHz-700 MHz..................................... 700 50
700 MHz-1 GHz....................................... 700 100
1 GHz-2 GHz......................................... 2000 200
2 GHz-4 GHz......................................... 3000 200
4 GHz-6 GHz......................................... 3000 200
6 GHz-8 GHz......................................... 1000 200
8 GHz-12 GHz........................................ 3000 300
12 GHz-18 GHz....................................... 2000 200
18 GHz-40 GHz....................................... 600 200
------------------------------------------------------------------------
The field strengths are expressed in terms of peak root-mean-square
(rms) values over the complete modulation period.
The threat levels identified above are the result of an FAA review
of existing studies on the subject of HIRF.
15. Operation Without Normal Electrical Power
This Special Condition was developed to address fly-by-wire
airplanes starting with the Airbus Model A330. As with earlier
airplanes, the Airbus A380-800 fly-by-wire control system requires a
continuous source of electrical power for the flight control system to
remain operable.
Section 25.1351(d), ``Operation without normal electrical power,''
requires safe operation in visual flight rules (VFR) weather conditions
for at least five minutes with inoperative normal power. This rule was
structured around a traditional design utilizing mechanical control
cables for flight control while the crew took time to sort out the
electrical failure, start the engine(s) if necessary, and re-establish
some of the electrical power generation capability.
To maintain the same level of safety as that associated with
traditional designs, the Model A380 design must not be time limited in
its operation, including being without the normal source of engine or
Auxiliary Power Unit (APU) generated electrical power. Service
experience has shown that the loss of all electrical power generated by
the airplane's engine generators or APU is not extremely improbable.
Thus, it must be demonstrated that the airplane can continue through
safe flight and landing--including steering and braking on the ground
for airplanes using steer/brake-by-wire--using its emergency electrical
power systems. These emergency electrical power systems must be able to
power loads that are essential for continued safe flight and landing.
Discussion of Comments
Notice of Proposed Special Conditions No. 25-04-05-SC for the
Airbus A380 airplane was published in the Federal Register on April 12,
2005 (70 FR 19015). The only commenter, the Boeing Company, submitted
comments on all proposed Special Conditions, except Special Condition
No. 12.
Boeing submitted comments in support of proposed Special Conditions
No. 1, 3, 4, 8, and 11. No change to those special conditions was
requested. In addition, Boeing submitted comments requesting a change
to proposed Special Conditions 2, 5, 6, 7, 9, 10, 12, 13, 14, and 15.
Those comments are discussed below.
Comments on Special Conditions No. 2. Interaction of Systems and
Structures
Requested change 1: The Boeing Company states that paragraph
c.(2)(d), Warning considerations, ``should be revised to use
nomenclature that is consistent with 14 CFR 25.1322 and, thus, less
onerous on system failure detection expectations.'' Specifically,
[[Page 18174]]
Boeing suggests using the text of the final version of the Load and
Dynamics Harmonization Working Group (LDHWG) report of January 2003
that was accepted by the Aviation Rulemaking Advisory Committee (ARAC).
FAA response: The FAA agrees, in part, with this comment and,
accordingly, has changed the sentence which states ``The flight crew
must be made aware of these failures before flight,'' to ``As far as
reasonably practicable, the flight crew must be made aware of these
failures before flight.'' The other changes suggested would not
substantively affect the Special Conditions and, therefore, were not
adopted. The FAA does not agree, however, that retaining the proposed
nomenclature makes the requirement more onerous.
Requested change 2: The Boeing Company says that proposed Special
Conditions No. 2, paragraph c (2)(e), Dispatch with known failure
conditions, ``should be revised to stay within the scope of Part 25.''
Boeing adds that the proposed Special Conditions ``is attempting to
require what is acceptable for [Minimum Equipment List] MEL dispatch
with system failures, which falls under part 121 requirements
(specifically 14 CFR 121.628). Dispatch considerations and intervals
should be determined in coordination with the Flight Operations
Evaluation Board (FOEB) in establishing the Master Minimum Equipment
List (MMEL).''
Specifically, Boeing objects to the fact that the proposed Special
Conditions ``excludes the consideration of the probability of
dispatching with known failures to be considered in the Time of
Occurrence loads conditions, described in paragraph c. (2)(c)(1) and
its Figure 1 (Factor of safety at the time of occurrence). This would
effectively preclude failure conditions that meet the no-single-failure
criterion and are almost, but not quite, extremely improbable without
this dispatch probability consideration.''
FAA response: The FAA does not agree that a certification standard
for what is acceptable when the airplane is dispatched with known
failure conditions is outside the scope of part 25. Acceptable dispatch
configurations for the airplane are essentially variations of the type
design and, as such, should not compromise the level of safety provided
by the airplane's certification basis. Section 121.628 does not contain
standards by which to judge the safety of MMEL dispatch configurations.
It is the certification basis for the airplane, including any special
conditions, that provides these standards. Limitations on acceptable
dispatch configurations are legitimate subjects of these standards, and
such limitations have been included previously on Special Conditions
pertaining to Interaction of Systems and Structures. Such limitations
may be necessary, depending on the severity of the potential
consequences of failure conditions that could occur following dispatch
under the MMEL.
In terms of the comment that the proposed Special Conditions would
``effectively preclude failure conditions that meet the no-single-
failure criterion * * * '' we agree that the Special Conditions should
be clearer about how the provisions of paragraph (c) and Figure 1
apply. We have revised the text of Special Conditions No. 2, paragraph
c (2)(e), accordingly.
Comments on Special Conditions No. 5. Dive Speed Definition
Requested change 1: The Boeing Company states that on the design
for the Boeing Model 777, a dive speed definition with a speed
protection system was the subject of an equivalent level of safety
finding. According to Boeing, ``since the Model A380 is similarly
pursuing relief from the Dive Speed Definition, it should also be
required to include bank angle protection features designed to failure
rates less than 10E-5 per flight hour in order to be consistent with
previous FAA positions.''
FAA response: The FAA does not agree. The A380 does not have the
same protective functions as the Boeing Model 777. In particular, it
does not have a similar bank angle protection feature. However, the
A380 has protective systems that compensate for a reduced speed margin.
The proposed Special Conditions specify maximum failure rates for these
protective systems which are consistent with the approach taken on the
Boeing 777. Accordingly, we have not changed the text of proposed
Special Conditions No. 5.
Requested change 2: The Boeing Company also suggests that the
maximum failure rate specified for the protective systems is stated
differently in the equivalent level of safety finding for the Boeing
Model 777 airplane and in the Special Conditions proposed for the A380.
Boeing says, ``For consistency of application and interpretation, the
FAA should revise the Special Conditions to require that each of the
A380 compensating features also meet the minimum 10E-5 failure rate
criterion.''
FAA response: The FAA does not agree. The A380 includes failure
annunciation features not included in the Boeing 777. The FAA
considered these annunciation features and follow-on pilot actions
defined in the airplane flight manual in determining adequate
requirements for maximum failure rate for the A380 protective systems.
We determined that a higher maximum failure rate (10E-3 per flight
hour) for such systems would provide adequate overall airplane level
protection. The FAA did not consider such annunciation features and
follow-on pilot actions during certification of the Boeing 777, because
such features were not presented to the FAA by the Boeing Company.
Nevertheless, the FAA considers the overall airplane level of
protection to be essentially the same in the two cases.
Comments on Special Conditions No. 6. Electronic Flight Control System:
Lateral-directional Stability, Longitudinal Stability, and Low Energy
Awareness
Requested change 1: The Boeing Company says that in the
certification programs for Airbus Models A330, A340, and A340-500/600,
the Special Conditions required demonstration of ``dynamic'' and
``static'' longitudinal stability and that the same requirement should
be added for consistency.
FAA response: The FAA does not agree. In past certification
programs on Airbus airplanes with electronic flight control systems, a
requirement to demonstrate dynamic stability was included in Special
Conditions, because the FAA initially thought that the requirement for
heavy damping of any short period oscillation, as contained in Sec.
25.181(a), might not be appropriate for the electronic flight control
system of Airbus airplanes. However, the FAA later learned that direct
compliance with Sec. 25.181 (a) could be demonstrated on Airbus
airplanes.
When Airbus initiated the certification process for the A380, the
FAA and the Joint Aviation Authorities (JAA) harmonized their
corresponding Special Conditions, including that pertaining to
Electronic Flight Control System-Longitudinal Stability. As a result of
the transition of authority from the JAA to the European Aviation
Safety Agency (EASA), EASA is now the certifying authority for the
Airbus A380 airplane. This harmonized A380 Special Conditions does not
include a dynamic requirement, because direct compliance with Sec.
25.181(a) will be demonstrated. Therefore, we have not revised the text
of the proposed Special Conditions.
Requested change 2: Boeing suggests that some of the qualifying
terms used are not defined, so that the Special
[[Page 18175]]
Conditions may not be applied consistently.
FAA response: The FAA agrees that--when we use words which have a
specific meaning in the context of a Special Conditions--we should
define or explain them. Therefore, we have revised the text of the
Special Conditions to add definitions of the terms ``suitable'' and
``adequate awareness.''
Comments on Special Conditions No. 7. Electronic Flight Control System:
Control Surface Awareness
Requested change: The Boeing Company comments that, ``The intent of
these Special Conditions is to provide suitable annunciation to the
flight crew when the flight control surfaces are close to their
authority limits without crew awareness.'' Boeing notes that ``in a
similar recent Issue Paper on the Boeing Model 787, the FAA references
autopilot back-drive in flight conditions described in these Special
Conditions. Without autopilot back-drive, control saturation is further
exacerbated.'' The company suggests that a crew procedure be required
when control saturation occurs along with Airplane Flight Manual (AFM)
instructions.
FAA response: The FAA does not agree. The Special Conditions for
indication of flight control position are relevant to electronic flight
control systems, regardless of whether or not the pilots' controls are
back-driven. While it is true that the differences in the designs may
affect the magnitude of the difference between control position and
surface position, the basic requirement for surface position awareness
applies to both design types. Both the A380 Special Conditions and the
787 Special Conditions issue paper noted by Boeing refer to the need
for a specific crew action. For both airplanes, the acceptability of
those crew actions will be determined as part of finding compliance
with their associated Special Conditions. However, the differences in
the designs do not warrant an additional, specific requirement for a
crew procedure based solely on the fact that the A380 control is not
back-driven.
The Boeing Company further requests that the statement ``without
being commanded by the crew or autopilot'' be included in the Special
Conditions. The FAA does not agree with this request, because the
suggested change would exclude the autopilot from the basic Special
Conditions requirement to provide an annunciation to the flight crew.
The autopilot drives the control surface without pilot input and,
therefore, could create flight conditions in which the control surface
deflection is approaching a limit without being commanded by the crew.
Accordingly, we have not changed the text of the proposed Special
Conditions.
Comments on Special Conditions No. 9. Flight Envelope Protection:
General Limiting Requirements
Requested change: The Boeing Company observes that Special
Conditions issued for earlier Airbus models that employ envelope
protection functions within the Electronic Flight Control System (EFCS)
have specifically addressed abnormal attitudes, while the proposed
Special Conditions for the Model A380 do not. Specifically, Boeing
suggests ``revising the proposed Special Conditions by adding a
paragraph to address abnormal attitudes and EFCS impact on recovery to
normal attitudes.''
FAA response: The FAA agrees that the paragraph addressing abnormal
attitudes should be included in the Special Conditions as in past
certification programs on Airbus airplanes. It was the FAA's intent to
cover this topic in other Special Conditions, in order to harmonize
with the approach used by the JAA. As a result of administrative
oversight, the FAA did not include this topic in other Special
Conditions, so it has been added to Special Condition No. 9. Since this
requirement has been included in multiple previous FAA Special
Conditions for Airbus airplanes without significant public comment, the
FAA has determined that it can be added to Special Condition No. 9
without further notice and comment.
Comments on Special Conditions No. 10. Flight Envelope Protection:
Normal Load Factor (G) Limiting
Requested change: The Boeing Company states that the text of these
Special Conditions differs from similar ones issued previously for
Airbus Models A320, A330, and A340, in that the phrase ``in the absence
of other limiting factors'' has been added as a condition of applying
the required action. Boeing suggests that, ``With this additional
phrase, the applicability of this Special Conditions is ambiguous; it
allows this Special Conditions essentially to be ignored when other
`limiting factors' are present.'' Therefore, Boeing recommends that the
phrase be either removed or explained.
FAA response: The phrase ``in the absence of other limiting
factors'' was added to the proposed Special Conditions to harmonize
with the JAA. The FAA does not agree that the phrase is ambiguous or
that it allows the Special Conditions to be ignored when other limiting
factors are present. It simply means that there are other limiting
factors, such as those discussed in the preamble, that would establish
the upper boundary for normal load factor and that the Special
Conditions are addressing only the lower boundary. Accordingly, we have
not revised the text of the proposed Special Conditions but have added
a sentence of explanation to the preamble.
Comment on Special Conditions No. 13. Flight Envelope Protection: High
Incidence Protection and Alpha-Floor Systems
Requested change 1: The Boeing Company recommends that we ``change
the procedure for determining minimum operating speeds, so that angle-
of-attack limiting envelope protection functions are active during the
maneuvers used to define the Reference Stall Speed.'' Boeing also
requests that paragraph c. (5)(g) specify that the high incidence
protection system should be ``operating normally'' instead of
``adjusted to a high enough incidence to allow full development of the
1g stall.''
FAA response: The meaning of the request is unclear, since it is
not the intent of paragraph c. (5) to determine either minimum
operating speeds or the reference stall speed. The FAA does not agree
with the request to revise the text. The intent of paragraph c. (5) is
to set the conditions for determining VCLMAX as defined in
paragraph c. (4). Without adjusting the high incidence protection
system angle, it would not be possible to achieve the 1g stall speed,
VCLMAX. VCLMAX is not a minimum operating speed
but rather a speed that depends on a specific test procedure and on the
stall characteristics of the airplane. The reference stall speed is
selected by the applicant, but it must be greater than or equal to
VCLMAX. Accordingly, we have not revised the text of the
proposed Special Conditions.
Requested change 2: The Boeing Company suggests that--to be
consistent with the criteria, intent, and philosophy of prior Issue
Papers and Special Conditions--certain changes be made to the proposed
Special Conditions. These changes pertain to (1) failure annunciation,
(2) prohibition of dispatch with the high incidence protection and
alpha floor systems inoperative, (3) additional demonstration for alpha
floor system inoperative, and (4) testing with system components set to
adverse tolerances limits.
FAA Response. (1) Failure Annunciation: The FAA does not agree that
annunciation of failure of the stall protection system and loss of
control
[[Page 18176]]
capability should be specified in these Special Conditions.
Annunciation of a system failure condition is covered in Sec.
25.1309(c). Paragraph 13(d)(2) of these Special Conditions states that
stall warning must be provided in accordance with Sec. 25.207
following failures of the high incidence protection system not shown to
be extremely improbable.
(2) No dispatch with system inoperative: As noted in the FAA
response to Boeing's comment on Special Condition No. 2, the FAA has
the authority, under part 25, to identify limitations to dispatch
configurations in the MMEL, when necessary for type certification.
However, in the case of Special Condition No. 13, we have determined
that specific limitations on dispatch following failures of the high
incidence protection and alpha floor protection systems are not needed
for type certification. The FAA Flight Operations Evaluation Board
should still determine the dispatch capability of the A380 relevant to
these two systems, as part of the their normal processes for
operational approvals.
(3) Additional demonstration for alpha floor system inoperative:
The FAA does not agree that--to satisfy the intent of paragraph d(2)--
the requirement should include the failure of the alpha floor system.
Paragraph d(2) refers to paragraphs b(1), (2), and (3), and states that
stall warning must be provided if these requirements are not met. The
alpha floor system is independent of the high incidence protection
system. If the alpha floor system fails, it should have no effect on
the function and requirements of the high incidence protection system
and should not invoke stall warning.
(4) Requirement to test with system components set to adverse
tolerance limits: The Boeing Company suggests that the Special
Conditions require that ``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.'' The FAA agrees that the above statement should be
included in these Special Conditions. However, as this statement also
pertains to production tolerances for the angle-of-attack protection
system, application to the Airbus A380 should include tolerances for
the angle-of-attack limits set for the high incidence protection system
as well as for the backup stall warning system. The FAA has revised the
text of the Special Conditions, accordingly.
Comments on Special Conditions No. 14. High Intensity Radiated Fields
(HIRF) Protection
Requested change: The Boeing Company says that the requirement for
``engineering validation of maintenance'' which has been included in
previous Special Conditions is not included and requests that it be
added.
FAA Response: ``Engineering validation of maintenance'' is a method
of compliance issue that is addressed in issue papers. It has not been
included in previously-published special conditions and is not
appropriate for Special Condition No. 14.
Comments on Special Condition No. 15. Operation Without Normal
Electrical Power
Requested change: The Boeing Company comments that, ``this proposed
Special Condition is attempting to advance safety standards through the
use of Special Conditions'' and that ``the current regulations,
Sec. Sec. 25.1351(d), 25.671(d) and 25.1309, considering the intended
operation of the airplane and its longest diversion, provide
appropriate and adequate safety standards.'' Boeing requests that the
proposed Special Conditions be replaced with information about
appropriate means of compliance.
FAA response: The FAA does not agree. The A380 design incorporates
electronic flight controls which are a new and novel feature not
envisioned when Sec. 25.1351(d) was promulgated. In addition, Sec.
25.1351(d) is inadequate, because it requires only 5 minutes of standby
power. The A380 would be incapable of continued safe flight and landing
with less than 5 minutes of standby power. Therefore, Special
Conditions that address operations without normal electrical power are
appropriate for the A380 fly-by-wire airplane, and we have not revised
the text of the proposed Special Conditions.
Clarification
In addition to changes made in responses to comments, the FAA has
revised the wording of one of the provisions of Special Conditions No.
13, Flight Envelope Protection: High Incidence Protection and Alpha-
floor Systems. The wording of paragraph j (1) has been slightly revised
to clarify the intent.
Applicability
As discussed above, these Special Conditions are applicable to the
Airbus A380-800 airplane. Should Airbus apply at a later date for a
change to the type certificate to include another model incorporating
the same novel or unusual design features, these Special Conditions
would apply to that model as well under the provisions of Sec. 21.101.
Conclusion
This action affects only certain novel or unusual design features
of the Airbus A380-800 airplane. It is not a rule of general
applicability, and it affects only the applicant that 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 to me by the
Administrator, the following Special Conditions are issued as part of
the type certification basis for the Airbus A380-800 airplane.
1. Dynamic Braking
In addition to the requirements of Sec. 25.493(d), the following
Special Conditions apply:
Loads arising from the sudden application of maximum braking effort
must be defined, taking into account the behavior of the braking
system. Failure conditions of the braking system must be analyzed in
accordance with the criteria specified in Special Conditions No. 2,
``Interaction of Systems and Structures.''
2. Interaction of Systems and Structures
In addition to the requirements of part 25, subparts C and D, the
following Special Conditions apply:
a. For airplanes equipped with systems that affect structural
performance--either directly or as a result of a failure or
malfunction--the influence of these systems and their failure
conditions must be taken into account when showing compliance with the
requirements of part 25, subparts C and D. Paragraph c. below must be
used to evaluate the structural performance of airplanes equipped with
these systems.
b. Unless shown to be extremely improbable, the airplane must be
designed to withstand any forced structural vibration resulting from
any failure, malfunction, or adverse condition in the flight control
system. These loads must be treated in
[[Page 18177]]
accordance with the requirements of paragraph a. above.
c. Interaction of Systems and Structures
(1) General: The following criteria must be used for showing
compliance with these Special Conditions and with Sec. 25.629 for
airplanes equipped with flight control systems, autopilots, stability
augmentation systems, load alleviation systems, flutter control
systems, and fuel management systems. If this paragraph is used for
other systems, it may be necessary to adapt the criteria to the
specific system.
(a) The criteria defined herein address only the direct structural
consequences of the system responses and performances. They cannot be
considered in isolation but should be included in the overall safety
evaluation of the airplane. These criteria may, in some instances,
duplicate standards already established for this evaluation. These
criteria are applicable only to structures whose failure could prevent
continued safe flight and landing. Specific criteria that define
acceptable limits on handling characteristics or stability requirements
when operating in the system degraded or inoperative modes are not
provided in this paragraph.
(b) Depending upon the specific characteristics of the airplane,
additional studies may be required that go beyond the criteria provided
in this paragraph in order to demonstrate the capability of the
airplane to meet other realistic conditions, such as alternative gust
or maneuver descriptions for an airplane equipped with a load
alleviation system.
(c) The following definitions are applicable to this paragraph.
Structural performance: Capability of the airplane to meet the
structural requirements of part 25.
Flight limitations: Limitations that can be applied to the airplane
flight conditions following an in-flight occurrence and that are
included in the flight manual (e.g., speed limitations and avoidance of
severe weather conditions).
Operational limitations: Limitations, including flight limitations,
that can be applied to the airplane operating conditions before
dispatch (e.g., fuel, payload, and Master Minimum Equipment List
limitations).
Probabilistic terms: The probabilistic terms (probable, improbable,
and extremely improbable) used in this Special 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, this Special Conditions applies only to
system failure conditions that affect the structural performance of the
airplane (e.g., system failure conditions that induce loads, change the
response of the airplane to inputs such as gusts or pilot actions, or
lower flutter margins).
(2) Effects of Systems on Structures.
(a) General. The following criteria will be used in determining the
influence of a system and its failure conditions on the airplane
structure.
(b) System fully operative. With the system fully operative, the
following apply:
(1) Limit loads must be derived in all normal operating
configurations of the system from all the limit conditions specified in
Subpart C, taking into account any special behavior of such a system or
associated functions or any effect on the structural performance of the
airplane that may occur up to the limit loads. In particular, any
significant non-linearity (rate of displacement of control surface,
thresholds or any other system non-linearities) must be accounted for
in a realistic or conservative way when deriving limit loads from limit
conditions.
(2) The airplane must meet the strength requirements of part 25
(Static strength, residual strength), using the specified factors to
derive ultimate loads from the limit loads defined above. The effect of
non-linearities must be investigated beyond limit conditions to ensure
that the behavior of the system presents no anomaly compared to the
behavior below limit conditions. However, conditions beyond limit
conditions need not be considered, when it can be shown that the
airplane has design features that will not allow it to exceed those
limit conditions.
(3) The airplane must meet the aeroelastic stability requirements
of Sec. 25.629.
(c) System in the failure condition. For any system failure
condition not shown to be extremely improbable, the following apply:
(1) At the time of occurrence. Starting from 1g level flight
conditions, a realistic scenario, including pilot corrective actions,
must be established to determine the loads occurring at the time of
failure and immediately after failure.
(i) For static strength substantiation, these loads multiplied by
an appropriate factor of safety that is related to the probability of
occurrence of the failure are ultimate loads to be considered for
design. The factor of safety (FS) is defined in Figure 1.
[GRAPHIC] [TIFF OMITTED] TR11AP06.009
[[Page 18178]]
(ii) For residual strength substantiation, the airplane must be
able to withstand two thirds of the ultimate loads defined in Paragraph
(c)(1)(i) of this section.
(iii) Freedom from aeroelastic instability must be shown up to the
speeds defined in Sec. 25.629(b)(2). For failure conditions that
result in speed increases beyond VC/MC, freedom
from aeroelastic instability must be shown to increased speeds, so that
the margins intended by Sec. 25.629(b)(2) are maintained.
(iv) Failures of the system that result in forced structural
vibrations (oscillatory failures) must not produce loads that could
result in detrimental deformation of primary structure.
(2) For the continuation of the flight. For the airplane in the
system failed state and considering any appropriate reconfiguration and
flight limitations, the following apply:
(i) The loads derived from the following conditions at speeds up to
Vc or the speed limitation prescribed for the remainder of the flight
must be determined:
(A) the limit symmetrical maneuvering conditions specified in Sec.
25.331 and in Sec. 25.345.
(B) the limit gust and turbulence conditions specified in Sec.
25.341 and in Sec. 25.345.
(C) the limit rolling conditions specified in Sec. 25.349 and the
limit unsymmetrical conditions specified in Sec. 25.367 and Sec.
25.427(b) and (c).
(D) the limit yaw maneuvering conditions specified in Sec. 25.351.
(E) the limit ground loading conditions specified in Sec. 25.473
and Sec. 25.491.
(ii) For static strength substantiation, each part of the structure
must be able to withstand the loads in Paragraph (2)(i) of this Special
Conditions 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] TR11AP06.010
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: If Pj is greater than 10-3 per
flight hour, then a 1.5 factor of safety must be applied to all
limit load conditions specified in Subpart C.
(iii) For residual strength substantiation, the airplane must be
able to withstand two thirds of the ultimate loads defined in Paragraph
(c)(2)(ii).
(iv) If the loads induced by the failure condition have a
significant effect on fatigue or damage tolerance, then their effects
must be taken into account.
(v) Freedom from aeroelastic instability must be shown up to a
speed determined from Figure 3. Flutter clearance speeds V' and V'' may
be based on the speed limitation specified for the remainder of the
flight, using the margins defined by Sec. 25.629(b).
[GRAPHIC] [TIFF OMITTED] TR11AP06.011
[[Page 18179]]
V' = Clearance speed as defined by Sec. 25.629(b)(2).
V'' = 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: If Pj is greater than 10-3 per
flight hour, then the flutter clearance speed must not be less than
V'
(vi) Freedom from aeroelastic instability must also be shown up to
V' in Figure 3 above for any probable system failure condition combined
with any damage required or selected for investigation by Sec.
25.571(b).
(3) Consideration of certain failure conditions may be required by
other sections of this Part, regardless of calculated system
reliability. Where analysis shows the probability of these failure
conditions to be less than 10-\9\, criteria other than those
specified in this paragraph may be used for structural substantiation
to show continued safe flight and landing.
(d) Warning considerations. For system failure detection and
warning, the following apply:
(1) The system must be checked for failure conditions, not
extremely improbable, that degrade the structural capability below the
level required by part 25 or significantly reduce the reliability of
the remaining system. As far as reasonably practicable, the flight crew
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
systems and where service history shows that inspections will provide
an adequate level of safety.
(2) The existence of any failure condition, not extremely
improbable, during flight that could significantly affect the
structural capability of the airplane and for which the associated
reduction in airworthiness can be minimized by suitable flight
limitations must be signaled to the flightcrew. For example, failure
conditions that result in a factor of safety between the airplane
strength and the loads of part 25, subpart C, below 1.25 or flutter
margins below V'' must be signaled to the crew during flight.
(e) Dispatch with known failure conditions. If the airplane is to
be dispatched in a known system failure condition that affects
structural performance or affects the reliability of the remaining
system to maintain structural performance, then the provisions of this
Special Conditions must be met, including the provisions of Paragraph
(b), for the dispatched condition and Paragraph (c) for subsequent
failures. Expected operational limitations may be taken into account in
establishing Pj as the probability of failure occurrence for
determining the safety margin in Figure 1. Flight limitations and
expected operational limitations may be taken into account in
establishing Qj as the combined probability of being in the dispatched
failure condition and the subsequent failure condition for the safety
margins in Figures 2 and 3. These limitations must be such that the
probability of being in this combined failure state and then
subsequently encountering limit load conditions is extremely
improbable. No reduction in these safety margins is allowed, if the
subsequent system failure rate is greater than 1E-3 per flight hour.
3. Limit Pilot Forces
In addition to the requirements of Sec. 25.397(c) the following
Special Conditions apply: The limit pilot forces are as follows:
a. For all components between and including the handle and its
control stops.
------------------------------------------------------------------------
Pitch Roll
------------------------------------------------------------------------
Nose up 200 lbf........................... Nose left 100 lbf.
Nose down 200 lbf......................... Nose right 100 lbf.
------------------------------------------------------------------------
b. For all other components of the side stick control assembly, but
excluding the internal components of the electrical sensor assemblies
to avoid damage as a result of an in-flight jam.
------------------------------------------------------------------------
Pitch Roll
------------------------------------------------------------------------
Nose up 125 lbf........................... Nose left 50 lbf.
Nose down 125 lbf......................... Nose right 50 lbf.
------------------------------------------------------------------------
4. Side Stick Controllers
In the absence of specific requirements for side stick controllers,
the following Special Conditions apply:
a. Pilot strength: In lieu of the ``strength of pilots'' limits
shown in Sec. 25.143(c) for pitch and roll and in lieu of the specific
pitch force requirements of Sec. Sec. 25.145(b) and 25.175(d), it must
be shown that the temporary and maximum prolonged force levels for the
side stick controllers are suitable for all expected operating
conditions and configurations, whether normal or non-normal.
b. Pilot control authority: The electronic side stick controller
coupling design must provide for corrective and/or overriding control
inputs by either pilot with no unsafe characteristics. Annunciation of
the controller status must be provided and must not be confusing to the
flight crew.
c. Pilot control: It must be shown by flight tests that the use of
side stick controllers does not produce unsuitable pilot-in-the-loop
control characteristics when considering precision path control/tasks
and turbulence. In addition, pitch and roll control force and
displacement sensitivity must be compatible, so that normal inputs on
one control axis will not cause significant unintentional inputs on the
other.
d. Autopilot quick-release control location: In lieu of compliance
with 25.1329(d), autopilot quick release (emergency) controls must be
on both side stick controllers. The quick release means must be located
so that it can readily and easily be used by the flight crew.
5. Dive Speed Definition
In lieu of the requirements of Sec. 25.335(b)(1)--if the flight
control system includes functions which act automatically to initiate
recovery before the end of the 20 second period specified in Sec.
25.335(b)(1)--the greater of the speeds resulting from the following
Special Conditions applies.
a. From an initial condition of stabilized flight at VC/ MC, the
airplane is upset so as to take up a new flight path 7.5 degrees below
the initial path. Control application, up to full authority, is made to
maintain this new flight path. Twenty seconds after initiating the
upset, manual recovery is made at a load factor of 1.5 g (0.5
acceleration increment) or such greater load factor that is
automatically applied by the system with the pilot's pitch control
neutral. The speed increase occurring in this maneuver may be
calculated, if reliable or conservative aerodynamic data is used.
Power, as specified in Sec. 25.175(b)(1)(iv), is assumed until
recovery is made, at which time power reduction and the use of pilot
controlled drag devices may be used.
b. From a speed below VC/MC with power to
maintain stabilized level flight at this speed, the airplane is upset
so as to accelerate through VC/MC at a flight
path 15 degrees below the initial path--or at the steepest nose down
attitude that the system will permit with full control authority if
less than 15 degrees.
[[Page 18180]]
Note: The pilot's controls may be in the neutral position after
reaching VC/MC and before recovery is
initiated.
c. Recovery may be initiated three seconds after operation of high
speed warning system by application of a load of 1.5g (0.5 acceleration
increment) or such greater load factor that is automatically applied by
the system with the pilot's pitch control neutral. Power may be reduced
simultaneously. All other means of decelerating the airplane, the use
of which is authorized up to the highest speed reached in the maneuver,
may be used. The interval between successive pilot actions must not be
less than one second.
d. The applicant must also demonstrate either that
(1) the speed margin, established as above, will not be exceeded in
inadvertent or gust induced upsets, resulting in initiation of the dive
from non-symmetric attitudes, or
(2) the airplane is protected by the flight control laws from
getting into non-symmetric upset conditions.
e. The probability of failure of the protective system that
mitigates for the reduced speed margin must be less than
10-\5\ per flight hour, except that the probability of
failure may be greater than 10-\5\, but not greater than
10-\3\, per flight hour, provided that:
(1) Failures of the system are annunciated to the pilots, and
(2) The flight manual instructions require the pilots to reduce the
speed of the airplane to a value that maintains a speed margin between
VMO and VD consistent with showing compliance
with 25.335(b) without the benefit of the system, and
(3) no dispatch of the airplane is allowed with the system
inoperative.
6. Electronic Flight Control System: Lateral-Directional and
Longitudinal Stability and Low Energy Awareness
In lieu of the requirements of Sec. Sec. 25.171, 25.173, 25.175,
and 25.177(c), the following Special Conditions apply:
a. The airplane must be shown to have suitable static lateral,
directional, and longitudinal stability in any condition normally
encountered in service, including the effects of atmospheric
disturbance. The showing of suitable static lateral, directional, and
longitudinal stability must be based on the airplane handling
qualities, including pilot workload and pilot compensation, for
specific test procedures during the flight test evaluations.
b. The airplane must provide adequate awareness to the pilot of a
low energy (low speed/low thrust/low height) state when fitted with
flight control laws presenting neutral longitudinal stability
significantly below the normal operating speeds. ``Adequate awareness''
means warning information must be provided to alert the crew of unsafe
operating conditions and to enable them to take appropriate corrective
action.
c. The static directional stability--as shown by the tendency to
recover from a skid with the rudder free--must be positive for any
landing gear and flap position and symmetrical power condition, at
speeds from 1.13 VS1g up to VFE, VLE,
or VFC/MFC (as appropriate).
d. In straight, steady sideslips (unaccelerated forward slips), the
rudder control movements and forces must be substantially proportional
to the angle of sideslip, and the factor of proportionality must be
between limits found necessary for safe operation throughout the range
of sideslip angles appropriate to the operation of the airplane. At
greater angles--up to the angle at which full rudder control is used or
a rudder pedal force of 180 pounds (81.72 kg) is obtained--the rudder
pedal forces may not reverse, and increased rudder deflection must
produce increased angles of sideslip. Unless the airplane has a
suitable sideslip indication, there must be enough bank and lateral
control deflection and force accompanying sideslipping to clearly
indicate any departure from steady, unyawed flight.
7. Electronic Flight Control System: Control Surface Awareness
In addition to the requirements of Sec. Sec. 25.143, 25.671 and
25.672, the following Special Conditions apply:
a. A suitable flight control position annunciation must be provided
to the crew in the following situation:
A flight condition exists in which--without being commanded by the
crew--control surfaces are coming so close to their limits that return
to normal flight and (or) continuation of safe flight requires a
specific crew action.
b. In lieu of control position annunciation, existing indications
to the crew may be used to prompt crew action, if they are found to be
adequate.
Note: The term ``suitable'' also indicates an appropriate
balance between nuisance and necessary operation.
8. Electronic Flight Control System: Flight Characteristics Compliance
Via the Handling Quantities Rating Method (HQRM)
a. Flight Characteristics Compliance Determination for EFCS Failure
Cases:
In lieu of compliance with Sec. 25.672(c), the HQRM contained in
Appendix 7 of AC 25-7A must be used for evaluation of EFCS
configurations resulting from single and multiple failures not shown to
be extremely improbable.
The handling qualities ratings are as follows:
(1) Satisfactory: Full performance criteria can be met with routine
pilot effort and attention.
(2) Adequate: Adequate for continued safe flight and landing; full
or specified reduced performance can be met, but with heightened pilot
effort and attention.
(3) Controllable: Inadequate for continued safe flight and landing,
but controllable for return to a safe flight condition, safe flight
envelope and/or reconfiguration, so that the handling qualities are at
least Adequate.
b. Handling qualities will be allowed to progressively degrade with
failure state, atmospheric disturbance level, and flight envelope, as
shown in Figure 12 of Appendix 7. Specifically, for probable failure
conditions within the normal flight envelope, the pilot-rated handling
qualities must be satisfactory in light atmospheric disturbance and
adequate in moderate atmospheric disturbance. The handling qualities
rating must not be less than adequate in light atmospheric disturbance
for improbable failures.
Note: AC 25-7A, Appendix 7 presents a method of compliance and
provides guidance for the following:
Minimum handling qualities rating requirements in
conjunction with atmospheric disturbance levels, flight envelopes,
and failure conditions (Figure 12),
Flight Envelope definition (Figures 5A, 6 and 7),
Atmospheric Disturbance Levels (Figure 5B),
Flight Control System Failure State (Figure 5C),
Combination Guidelines (Figures 5D, 9 and 10), and
General flight task list, from which appropriate
specific tasks can be selected or developed (Figure 11).
9. Flight Envelope Protection
(a) General Limiting Requirements. (1) Onset characteristics of
each envelope protection feature must be smooth, appropriate to the
phase of flight and type of maneuver, and not in conflict with the
ability of the pilot to satisfactorily change the airplane flight path,
speed, or attitude, as needed.
(2) Limit values of protected flight parameters (and if applicable,
associated warning thresholds) must be compatible with the following:
(a) Airplane structural limits,
[[Page 18181]]
(b) Required safe and controllable maneuvering of the airplane, and
(c) Margins to critical conditions. Dynamic maneuvering, airframe
and system tolerances (both manufacturing and in-service), and non-
steady atmospheric conditions--in any appropriate combination and phase
of flight--must not result in a limited flight parameter beyond the
nominal design limit value that would cause unsafe flight
characteristics.
(3) The airplane must be responsive to intentional dynamic
maneuvering to within a suitable range of the parameter limit. Dynamic
characteristics, such as damping and overshoot, must also be
appropriate for the flight maneuver and limit parameter in question.
(4) When simultaneous envelope limiting is engaged, adverse
coupling or adverse priority must not result.
b. Failure States: EFCS failures, including sensor failures, must
not result in a condition where a parameter is limited to such a
reduced value that safe and controllable maneuvering is no longer
available. The crew must be alerted by suitable means, if any change in
envelope limiting or maneuverability is produced by single or multiple
failures of the EFCS not shown to be extremely improbable.
c. Abnormal Attitudes: In case of abnormal attitude or excursion of
any other flight parameters outside the protected boundaries, the
operation of the EFCS, including the automatic protection functions,
must not hinder airplane recovery.
10. Flight Envelope Protection: Normal Load Factor (g) Limiting
In addition to the requirements of 25.143(a)--and in the absence of
other limiting factors--the following Special Conditions apply:
a. The positive limiting load factor must not be less than:
(1) 2.5g for the EFCS normal state.
(2) 2.0g for the EFCS normal state with the high lift devices
extended.
b. The negative limiting load factor must be equal to or more
negative than:
(1) Minus 1.0g for the EFCS normal state.
(2) 0.0g for the EFCS normal state with high lift devices extended.
Note: This Special Condition does not impose an upper bound for
the normal load factor limit, nor does it require that the limit
exist. If the limit is set at a value beyond the structural design
limit maneuvering load factor ``n,'' indicated in Sec. 25.333(b)
and 25.337(b) and (c), there should be a very positive tactile feel
built into the controller and obvious to the pilot that serves as a
deterrent to inadvertently exceeding the structural limit.
11. Flight Envelope Protection: High Speed Limiting
In addition to Sec. 25.143, the following Special Condition
applies:
Operation of the high speed limiter during all routine and descent
procedure flight must not impede normal attainment of speeds up to the
overspeed warning.
12. Flight Envelope Protection: Pitch And Roll Limiting
In addition to Sec. 25.143, the following Special Conditions
apply:
a. The pitch limiting function must not impede normal maneuvering
for pitch angles up to the maximum required for normal maneuvering--
including a normal all-engines operating takeoff plus a suitable margin
to allow for satisfactory speed control.
b. The pitch and roll limiting functions must not restrict or
prevent attaining roll angles up to 65 degrees or pitch attitudes
necessary for emergency maneuvering. Spiral stability, which is
introduced above 33 degrees roll angle, must not require excessive
pilot strength to achieve roll angles up to 65 degrees.
13. Flight Envelope Protection: High Incidence Protection And Alpha-
floor Systems
a. Definitions. For the purpose of this Special Condition, the
following definitions apply:
High Incidence Protection System A system that operates directly
and automatically on the airplane's flying controls to limit the
maximum angle of attack that can be attained to a value below that at
which an aerodynamic stall would occur.
Alpha-Floor System. A system that automatically increases thrust on
the operating engines when the angle of attack increases through a
particular value.
Alpha Limit. The maximum angle of attack at which the airplane
stabilizes with the high incidence protection system operating and the
longitudinal control held on its aft stop.
Vmin. The minimum steady flight speed is the stabilized,
calibrated airspeed obtained when the airplane is decelerated at an
entry rate not exceeding 1 knot per second, until the longitudinal
pilot control is on its stop with the high incidence protection system
operating.
Vmin1g Vmin corrected to 1g conditions. It is
the minimum calibrated airspeed at which the airplane can develop a
lift force normal to the flight path and equal to its weight when at an
angle of attack not greater than that determined for Vmin.
b. Capability and Reliability of the High Incidence Protection
System: (1) It must not be possible to encounter a stall during pilot
induced maneuvers, and handling characteristics must be acceptable, as
required by paragraphs e and f below, entitled High Incidence Handling
Demonstrations and High Incidence Handling Characteristics
respectively.
(2) The airplane must be protected against stalling due to the
effects of windshears and gusts at low speeds, as required by paragraph
g below, entitled Atmospheric Disturbances.
(3) The ability of the high incidence protection system to
accommodate any reduction in stalling incidence resulting from residual
ice must be verified.
(4) The reliability of the system and the effects of failures must
be acceptable, in accordance with Sec. 25.1309 and Advisory Circular
25.1309-1A, System Design and Analysis.
(5) The high incidence protection system must not impede normal
maneuvering for pitch angles up to the maximum required for normal
maneuvering, including a normal all-engines operating takeoff plus a
suitable margin to allow for satisfactory speed control.
c. Minimum Steady Flight Speed and Reference Stall Speed: In lieu
of the requirements of Sec. 25.103, the following Special Conditions
apply:
(1) Vmin. The minimum steady flight speed, for the
airplane configuration under consideration and with the high incidence
protection system operating, is the final stabilized calibrated
airspeed obtained when the airplane is decelerated at an entry rate not
exceeding 1 knot per second until the longitudinal pilot control is on
its stop.
(2) The minimum steady flight speed, Vmin, must be
determined with:
(a) The high incidence protection system operating normally.
(b) Idle thrust.
(c) Alpha-floor system inhibited.
(d) All combinations of flap settings and landing gear positions.
(e) The weight used when VSR is being used as a factor
to determine compliance with a required performance standard.
(f) The most unfavorable center of gravity allowable, and
(g) The airplane trimmed for straight flight at a speed achievable
by the automatic trim system.
(3) Vmin1g is Vmin corrected to 1g
conditions. Vmin1g is the minimum calibrated airspeed at
which the airplane can develop a lift force normal to the flight path
and equal to its weight when at an angle of attack not greater
[[Page 18182]]
than that determined for Vmin. Vmin1g is defined
as follows:
[GRAPHIC] [TIFF OMITTED] TR11AP06.025
where n z w = load factor normal to the flight path at
Vmin
(4) The Reference Stall Speed, VSR, is a calibrated
airspeed selected by the applicant. VSR may not be less than
the 1g stall speed. VSR is expressed as:
[GRAPHIC] [TIFF OMITTED] TR11AP06.026
where
VCLMAX = Calibrated airspeed obtained when the load factor-
corrected lift coefficient
[GRAPHIC] [TIFF OMITTED] TR11AP06.035
is first a maximum during the maneuver prescribed in Paragraph (5)(h)
of this Special Conditions.
nzw = Load factor normal to the flight path at
VCLMAX
W = Airplane gross weight
S = Aerodynamic reference wing area, and
q = Dynamic pressure.
(5) VCLMAX must be determined with the following
conditions:
(a) Engines idling or--if that resultant thrust causes an
appreciable decrease in stall speed--not more than zero thrust at the
stall speed
(b) The airplane in other respects, such as flaps and landing gear,
in the condition existing in the test or performance standard in which
VSR is being used.
(c) The weight used when VSR is being used as a factor
to determine compliance with a required performance standard.
(d) The center of gravity position that results in the highest
value of reference stall speed.
(e) The airplane trimmed for straight flight at a speed achievable
by the automatic trim system, but not less than 1.13 VSR and
not greater than 1.3 VSR.
(f) The alpha-floor system inhibited.
(g) The high incidence protection system adjusted to a high enough
incidence to allow full development of the 1g stall.
(h) Starting from the stabilized trim condition, apply the
longitudinal control to decelerate the airplane so that the speed
reduction does not exceed one knot per second.
(6) The flight characteristics at the angle of attack for
CLMAX must be suitable in the traditional sense at FWD and
AFT CG in straight and turning flight at IDLE power. Although for a
normal production EFCS and steady full aft stick this angle of attack
for CLMAX cannot be achieved, the angle of attack can be
obtained momentarily under dynamic circumstances and deliberately in a
steady state sense with some EFCS failure conditions.
d. Stall Warning. (1) Normal Operation. If the conditions of
Paragraph b, Capability and Reliability of the High Incidence
Protection System, are satisfied, a level of safety equivalent to that
intended by Sec. 25.207, Stall Warning, must be considered to have
been met without provision of an additional, unique warning device.
(2) Failure Cases. Following failures of the high incidence
protection system not shown to be extremely improbable, if the system
no longer satisfies Paragraph b, Capability and Reliability of the High
Incidence Protection System, parts (1), (2), and (3), stall warning
must be provided in accordance with Sec. 25.207. The stall warning
should prevent inadvertent stall under the following conditions:
(a) Power off straight stall approaches to a speed 5 percent below
the warning onset.
(b) Turning flight stall approaches at entry rates up to 3 knots
per second when recovery is initiated not less than one second after
the warning onset.
Note: ``Unless angle of attack (AOA) protection system (high
incidence 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 the high incidence protection system, stall warning and
stall identification should be set at the low AOA tolerance limit.
High AOA tolerance limits should be used for characteristics
evaluations.''
e. High Incidence Handling Demonstrations. In lieu of the
requirements of Sec. 25.201, the following Special Conditions apply:
Maneuvers to the limit of the longitudinal control in the nose up
direction must be demonstrated in straight flight and in 30 degree
banked turns under the following conditions:
(1) The high incidence protection system operating normally.
(2) Initial power condition of:
(a) Power off.
(b) The power necessary to maintain level flight at 1.5
VSR1, where VSR1 is the reference stall speed
with the flaps in the approach position, the landing gear retracted,
and the maximum landing weight. The flap position to be used to
determine this power setting is that position in which the stall speed,
VSR1, does not exceed 110% of the stall speed,
VSR0, with the flaps in the most extended landing position.
(3) Alpha-floor system operating normally, unless more severe
conditions are achieved with alpha-floor inhibited.
(4) Flaps, landing gear and deceleration devices in any likely
combination of positions.
(5) Representative weights within the range for which certification
is requested, and
(6) The airplane trimmed for straight flight at a speed achievable
by the automatic trim system.
f. High Incidence Handling Characteristics. In lieu of the
requirements of Sec. 25.203, the following Special Conditions apply:
(1) In demonstrating the handling characteristics specified in
paragraphs (2), (3), (4), and (5) below, the following procedures must
be used:
(a) 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.
(b) 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.
(c) 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.
(2) Throughout maneuvers with a rate of deceleration of not more
than 1 knot per second, both in straight flight and in 30 degree banked
turns, the airplane's characteristics must be as follows:
(a) There must not be any abnormal airplane nose-up pitching.
(b) There must not be any uncommanded nose-down pitching that would
be indicative of stall. However, reasonable attitude changes associated
with stabilizing the incidence at alpha limit as the longitudinal
control reaches the stop would be acceptable. Any reduction of pitch
attitude associated with stabilizing the incidence at the alpha limit
should be achieved smoothly and at a low pitch rate, such that it is
not likely to be mistaken for natural stall identification.
(c) There must not be any uncommanded lateral or directional
motion, and the pilot must retain good lateral and directional control
by conventional use of the cockpit controllers throughout the maneuver.
(d) The airplane must not exhibit buffeting of a magnitude and
severity
[[Page 18183]]
that would act as a deterrent to completing the maneuver.
(3) 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.
(4) It must always be possible to reduce incidence by conventional
use of the controller.
(5) The rate at which the airplane can be maneuvered from trim
speeds associated with scheduled operating speeds, such as
V2 and VREF, up to alpha-limit must not be unduly
damped or significantly slower than can be achieved on conventionally
controlled transport airplanes.
g. Atmospheric Disturbances. Operation of the high incidence
protection system and the alpha-floor system must not adversely affect
aircraft control during expected levels of atmospheric disturbances or
impede the application of recovery procedures in case of windshear.
Simulator tests and analysis may be used to evaluate such conditions
but must be validated by limited flight testing to confirm handling
qualities at critical loading conditions.
h. Alpha-floor. The alpha-floor setting must be such that the
aircraft can be flown at normal landing operational speed and
maneuvered up to bank angles consistent with the flight phase,
including the maneuver capabilities specified in 25.143(g), without
triggering alpha-floor. In addition, there must be no alpha-floor
triggering, unless appropriate, when the airplane is flown in usual
operational maneuvers and in turbulence.
i. Proof of Compliance: In addition to the requirements of Sec.
25.21, the following Special Conditions apply:
The flying qualities must be evaluated at the most unfavorable
center of gravity position.
j. Longitudinal Control: (1) In lieu of the requirements of Sec.
25.145(a) and 25.145(a)(1), the following Special Conditions apply:
It must be possible--at any point between the trim speed for
straight flight and Vmin--to pitch the nose downward, so
that the acceleration to this selected trim speed is prompt, with:
The airplane trimmed for straight flight at the speed achievable by
the automatic trim system and at the most unfavorable center of
gravity;
(2) In lieu of the requirements of Sec. 25.145(b)(6), the
following Special Conditions apply:
With power off, flaps extended and the airplane trimmed at 1.3
VSR1, obtain and maintain airspeeds between Vmin
and either 1.6 VSR1 or VFE, whichever is lower.
k. Airspeed Indicating System: (1) In lieu of the requirements of
subsection 25.1323(c)(1), the following Special Conditions apply:
VMO to Vmin with the flaps retracted.
(2) In lieu of the requirements of subsection 25.1323(c)(2), the
following Special Conditions apply:
Vmin to VFE with flaps in the landing position.
14. High Intensity Radiated Fields (HIRF) Protection
a. Protection from Unwanted Effects of High-intensity Radiated
Fields. Each electrical and electronic system which performs critical
functions must be designed and installed to ensure that the operation
and operational capabilities of these systems to perform critical
functions are not adversely affected when the airplane is exposed to
high intensity radiated fields external to the airplane.
b. For the purposes of this Special Conditions, the following
definition applies: Critical Functions: Functions whose failure would
contribute to or cause a failure condition which would prevent the
continued safe flight and landing of the airplane.
15. Operation Without Normal Electrical Power
In lieu of the requirements of Sec. 25.1351(d), the following
Special Condition applies:
It must be demonstrated by test or combination of test and analysis
that the airplane can continue safe flight and landing with inoperative
normal engine and APU generator electrical power (i.e., electrical
power sources, excluding the battery and any other standby electrical
sources). The airplane operation should be considered at the critical
phase of flight and include the ability to restart the engines and
maintain flight for the maximum diversion time capability being
certified.
Issued in Renton, Washington, on March 30, 2006.
Ali Bahrami,
Manager, Transport Airplane Directorate, Aircraft Certification
Service.
[FR Doc. 06-3359 Filed 4-10-06; 8:45 am]
BILLING CODE 4910-13-P