[Federal Register: July 30, 2007 (Volume 72, Number 145)]
[Rules and Regulations]
[Page 41428-41433]
From the Federal Register Online via GPO Access [wais.access.gpo.gov]
[DOCID:fr30jy07-4]
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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 25
[Docket No. NM362 Special Conditions No. 25-354-SC]
Special Conditions: Boeing Model 787-8 Airplane; Interaction of
Systems and Structures, Electronic Flight Control System-Control
Surface Awareness, High Intensity Radiated Fields (HIRF) Protection,
Limit Engine Torque Loads for Sudden Engine Stoppage, and Design Roll
Maneuver Requirement
AGENCY: Federal Aviation Administration (FAA), DOT.
ACTION: Final special conditions.
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SUMMARY: These special conditions are issued for the Boeing Model 787-8
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 electronic flight control systems and high bypass engines.
These special conditions also pertain to the effects of such novel or
unusual design features, such as effects on the structural performance
of the airplane. Finally, these special conditions pertain to effects
of certain conditions on these novel or unusual design features, such
as the effects of high intensity radiated fields (HIRF). The applicable
airworthiness regulations do not contain adequate or appropriate safety
standards for these design features. These special conditions contain
the additional safety standards that the Administrator considers
necessary to establish a level of safety equivalent to that established
by the existing standards. Additional special conditions will be issued
for other novel or unusual design features of the Boeing Model 787-8
airplanes.
DATES: Effective Date: August 29, 2007.
FOR FURTHER INFORMATION CONTACT: Meghan Gordon, FAA, Standardization
Branch, ANM-113, Transport Airplane Directorate, Aircraft Certification
Service, 1601 Lind Avenue SW., Renton, Washington 98057-3356; telephone
(425) 227-2138; facsimile (425) 227-1149.
SUPPLEMENTARY INFORMATION:
Background
On March 28, 2003, Boeing applied for an FAA type certificate for
its new Boeing Model 787-8 passenger airplane. The Boeing Model 787-8
airplane will be an all-new, two-engine jet transport airplane with a
two-aisle cabin. The maximum takeoff weight will be 476,000 pounds,
with a maximum passenger count of 381 passengers.
Type Certification Basis
Under provisions of 14 Code of Federal Regulations (CFR) 21.17,
Boeing must show that Boeing Model 787-8 airplanes (hereafter referred
to as ``787'') meet the applicable provisions of 14 CFR part 25, as
amended by Amendments 25-1 through 25-117, except Sec. Sec. 25.809(a)
and 25.812, which will remain at Amendment 25-115. If the Administrator
finds that the applicable airworthiness regulations do not contain
adequate or appropriate safety standards for the 787 because of a novel
or unusual design feature, special conditions are prescribed under
provisions of 14 CFR 21.16.
In addition to the applicable airworthiness regulations and special
conditions, the 787 must comply with the fuel vent and exhaust emission
requirements of 14 CFR part 34 and the noise certification requirements
of part 36. In addition, the FAA must issue a finding of regulatory
adequacy pursuant to section 611 of Public Law 92-574, the ``Noise
Control Act of 1972''.
The FAA issues special conditions, as defined in Sec. 11.19, under
Sec. 11.38 and they become part of the type certification basis under
Sec. 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 or similar
novel or unusual design feature, the special conditions would also
apply to the other model under Sec. 21.101.
Discussion of Novel or Unusual Design Features
The 787 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 the 787 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.
Most of these special conditions are identical or nearly identical
to those previously required for type certification of the Model 777
series airplanes.
Most of these 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 the case of some of these
requirements, 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 787 in the near future.
1. Interaction of Systems and Structures
The 787 is equipped with systems that affect the airplane's
structural performance, either directly or as a result of failure or
malfunction. That is, the airplane's systems affect how it responds in
maneuver and gust conditions, and thereby affect its structural
capability. These systems may also affect the aeroelastic stability of
the airplane. Such systems represent a novel and unusual feature when
compared to the technology envisioned in the current airworthiness
standards. Special conditions are needed to require consideration of
the effects of systems on the structural capability and aeroelastic
stability of the airplane, both in the normal and in the failed state.
These special conditions require that the airplane meet the
structural requirements of subparts C and D of 14 CFR part 25 when the
airplane systems
[[Page 41429]]
are fully operative. The special conditions also require that the
airplane meet these requirements considering failure conditions. In
some cases, reduced margins are allowed for failure conditions based on
system reliability.
2. 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
787, the pilot is not aware of the actual surface deflection position
during flight maneuvers. This feature of this design is novel and
unusual when compared to the state of technology envisioned in the
airworthiness standards for transport category airplanes. These special
conditions are meant to 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.
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 that would cause loss of control or other unsafe
handling or performance situations.
These special conditions require that suitable annunciation be
provided to the flightcrew when a flight condition exists in which
nearly full control surface deflection occurs. Suitability of such an
annunciation must take into account that some pilot-demanded maneuvers,
such as a rapid roll, are necessarily associated with intended full or
nearly full control surface deflection. Simple alerting systems which
would function in both intended and unexpected control-limiting
situations must be properly balanced between providing needed crew
awareness and avoiding nuisance warnings.
3. High Intensity Radiated Fields (HIRF) Protection
The 787 will use 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. Increased power levels from radio
frequency transmitter and use of sensitive avionics/electronics and
electrical systems to command and control the airplane have made it
necessary to provide adequate protection.
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 787. These special conditions require
that avionics/electronics and electrical systems that perform critical
functions be designed and installed to preclude component damage and
interruption of function because of HIRF.
High-power radio frequency transmitters for radio, radar,
television, and satellite communications can adversely affect
operations of airplane electrical and electronic systems. Therefore,
immunity of critical avionics/electronics and electrical systems to
HIRF must be established. Based on surveys and analysis of existing
HIRF emitters, adequate protection from HIRF exists if airplane system
immunity is demonstrated when exposed to the HIRF environments in
either paragraph (a) OR (b) below:
(a) A minimum environment of 100 volts rms (root-mean-square) per
meter electric field strength from 10 KHz to 18 GHz.
(1) System elements and their associated wiring harnesses must be
exposed to this environment without benefit of airframe shielding.
(2) Demonstration of this level of protection is established
through system tests and analysis.
(b) An environment external to the airframe of the field strengths
shown in the table below for the frequency ranges indicated. Immunity
to both peak and average field strength components from the table must
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
------------------------------------------------------------------------
Field strengths are expressed in terms of peak root-mean-square (rms)
values over the complete modulation period.
The environment levels identified above are the result of an FAA
review of existing studies on the subject of HIRF and of the work of
the Electromagnetic Effects Harmonization Working Group of ARAC.
4. Limit Engine Torque Loads for Sudden Engine Stoppage
The 787 will have high-bypass engines with a chord-swept fan 112
inches in diameter. Engines of this size were not envisioned when Sec.
25.361, pertaining to loads imposed by engine seizure, was adopted in
1965. Worst case engine seizure events become increasingly more severe
with increasing engine size because of the higher inertia of the
rotating components.
Section 25.361(b)(1) requires that for turbine engine
installations, the engine mounts and supporting structures must be
designed to withstand a ``limit engine torque load imposed by sudden
engine stoppage due to malfunction or structural failure.'' Limit loads
are expected to occur about once in the lifetime of any airplane.
Section 25.306 requires that supporting structures be able to support
limit loads without detrimental permanent deformation, meaning that
supporting structures should remain serviceable after a limit load
event.
Since adoption of Sec. 25.361(b)(1), the size, configuration, and
failure modes of jet engines have changed considerably. Current engines
are much larger and are designed with large bypass fans. In the event
of a structural failure, these engines are capable of producing much
higher transient loads on the engine mounts and supporting structures.
As a result, modern high bypass engines are subject to certain
rare-but-severe engine seizure events. Service history shows that such
events occur far less frequently than limit load events. Although it is
important for the airplane to be able to support such rare loads safely
without failure, it is unrealistic to expect that no permanent
deformation will occur.
Given this situation, ARAC has proposed a design standard for
today's large engines. For the commonly-occurring deceleration events,
the proposed standard requires engine mounts and structures to support
maximum torques without detrimental permanent deformation. For the
rare-but-severe engine seizure events such as loss of any fan,
compressor, or turbine blade, the proposed standard requires engine
mounts and structures to support maximum torques without failure, but
allows for some deformation in the structure.
[[Page 41430]]
The FAA concludes that modern large engines, including those on the
787, are novel and unusual compared to those envisioned when Sec.
25.361(b)(1) was adopted and thus warrant special conditions. These
special conditions contain design criteria recommended by ARAC.
5. Design Roll Maneuver Requirement
The 787 is equipped with an electronic flight control system that
provides control of the aircraft through pilot inputs to the flight
computer. Current part 25 airworthiness regulations account for
``control laws,'' for which aileron deflection is proportional to
control stick deflection. They do not address any nonlinearities \1\ or
other effects on aileron actuation that may be caused by electronic
flight controls. Therefore, the FAA considers the flight control system
to be a novel and unusual feature compared to those envisioned when
current regulations were adopted. Since this type of system may affect
flight loads, and therefore the structural capability of the airplane,
special conditions are needed to address these effects.
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\1\ A nonlinearity is a situation where output does not change
in the same proportion as input.
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These special conditions differ from current requirements in that
they require that the roll maneuver result from defined movements of
the cockpit roll control as opposed to defined aileron deflections.
Also, these special conditions require an additional load condition at
design maneuvering speed (VA), in which the cockpit roll
control is returned to neutral following the initial roll input.
These special conditions differ from similar special conditions
applied to previous designs. These special conditions are limited to
the roll axis only, whereas previous special conditions also included
pitch and yaw axes. Special conditions are no longer needed for the yaw
axis because Sec. 25.351 was revised at Amendment 25-91 to take into
account effects of an electronic flight control system. No special
conditions are needed for the pitch axis because the applicant's
proposed method for the pitch maneuver takes into account effects of an
electronic flight control system.
Discussion of Comments
Notice of Proposed Special Conditions No. 25-06-15-SC for the 787
was published in the Federal Register on March 12, 2007 (72 FR 10941).
Only one comment was received and it addressed proposed Special
Conditions No. 5.
Comment on Special Conditions No. 5. Design Roll Maneuver Requirement
Requested change: The commenter, an individual, stated that the
paragraph dealing with Sec. 25.349(a) in the proposed special
conditions is a little confusing. Paragraphs (c) and (d) of the
proposed special conditions both refer to ``paragraph (2)''. But there
are no numbered paragraphs in proposed Special Conditions No. 5. The
commenter thought that the reference was to paragraph (2) of Sec.
25.349(a), but since Sec. 25.349(a) is superseded by the special
conditions, the commenter suggested that this may cause confusion.
FAA response: The reference to paragraph (2) in the proposed
special conditions was an error and we thank the commenter for pointing
it out. The reference should have been ``paragraph (b).'' We have
revised the final special conditions accordingly. Otherwise, all
special conditions are adopted as proposed.
Applicability
As discussed above, these special conditions are applicable to the
787. Should Boeing apply at a later date for a change to the type
certificate to include another model on the same type certificate
incorporating the same novel or unusual design features, these special
conditions would apply to that model as well.
Conclusion
This action affects only certain novel or unusual design features
of the 787. It is not a rule of general applicability.
List of Subjects in 14 CFR Part 25
Aircraft, Aviation safety, Reporting and recordkeeping
requirements.
0
The authority citation for these special conditions is as follows:
Authority: 49 U.S.C. 106(g), 40113, 44701, 44702, 44704.
The Special Conditions
0
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 Boeing Model 787-8 airplane.
1. Interaction of Systems and Structures
The Boeing Model 787-8 airplane is equipped with systems which
affect the airplane's structural performance either directly or as a
result of failure or malfunction. The influence of these systems and
their failure conditions must be taken into account when showing
compliance with requirements of subparts C and D of part 25 of Title 14
of the Code of Federal Regulations. The following criteria must be used
for showing compliance with these special conditions for airplanes
equipped with flight control systems, autopilots, stability
augmentation systems, load alleviation systems, flutter control
systems, fuel management systems, and other systems that either
directly or as a result of failure or malfunction affect structural
performance. If these special conditions are used for other systems, it
may be necessary to adapt the criteria to the specific system.
(a) The criteria defined here address only direct structural
consequences of system responses and performances. They cannot be
considered in isolation but should be included in the overall safety
evaluation of the airplane. They may in some instances duplicate
standards already established for this evaluation. These criteria are
only applicable to structure whose failure could prevent continued safe
flight and landing. Specific criteria defining acceptable limits on
handling characteristics or stability requirements when operating in
the system degraded or inoperative mode are not provided in these
special conditions.
(b) Depending on the specific characteristics of the airplane,
additional studies may be required that go beyond the criteria provided
in these special conditions in order to demonstrate capability of the
airplane to meet other realistic conditions such as alternative gust
conditions or maneuvers for an airplane equipped with a load
alleviation system.
(c) The following definitions are applicable to these special
conditions.
(1) Structural performance: Capability of the airplane to meet the
structural requirements of part 25.
(2) Flight limitations: Limitations that can be applied to the
airplane flight conditions following an in-flight failure occurrence
and that are included in the flight manual (speed limitations or
avoidance of severe weather conditions, for example).
(3) Operational limitations: Limitations, including flight
limitations, that can be applied to the airplane operating conditions
before dispatch (fuel, payload, and master minimum equipment list
limitations, for example).
(4) Probabilistic terms: Terms (probable, improbable, extremely
improbable) used in these special conditions which are the same as
those probabilistic terms used in Sec. 25.1309.
(5) Failure condition: Term that is the same as that used in Sec.
25.1309. The term failure condition in these special conditions,
however, applies only to system failure conditions that affect
[[Page 41431]]
structural performance of the airplane. Examples are system failure
conditions that induce loads, change the response of the airplane to
inputs such as gusts or pilot actions, or lower flutter margins.
Note: Although failure annunciation system reliability must be
included in probability calculations for paragraph (f) of these
special conditions, there is no specific reliability requirement for
the annunciation system required in paragraph (g) of the special
conditions.
(d) General. The following criteria will be used in determining the
influence of a system and its failure conditions on the airplane
structure.
(e) 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 of 14 CFR part 25 (or used in lieu of those 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 degree of nonlinearity in rate of displacement of control
surface or thresholds, or any other system nonlinearities, 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 for
static strength and residual strength, using the specified factors to
derive ultimate loads from the limit loads defined above. The effect of
nonlinearities must be investigated beyond limit conditions to ensure
the behavior of the system presents no anomaly compared to the behavior
below limit conditions. However, conditions beyond limit conditions
need not be considered if the applicant demonstrates 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.
(f) System in the failure condition. For any system failure
condition not shown to be extremely improbable, the following apply:
(1) Establishing loads at the time of failure. Starting from 1-g
level flight conditions, a realistic scenario, including pilot
corrective actions, must be established to determine 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 related to 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] TR30JY07.000
(ii) For residual strength substantiation, the airplane must be
able to withstand two thirds of the ultimate loads defined in
subparagraph (f)(1)(i) of these special conditions. for pressurized
cabins, these loads must be combined with the normal operating
differential pressure.
(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 speeds beyond design cruise speed or design cruise mach
number (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) Establishing loads in the system failed state for the
continuation of the flight. For the continuation of flight of the
airplane in the system failed state and considering any appropriate
reconfiguration and flight limitations, the following apply:
(i) Loads derived from the following conditions (or used in lieu of
the following conditions) at speeds up to Vc/Mc,
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 Sec. 25.345.
(B) The limit gust and turbulence conditions specified in Sec.
25.341 and 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 (f)(2)(i) of these
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.
Figure 2
Factor of Safety For Continuation of Flight
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--Structure, of 14 CFR part 25.
[[Page 41432]]
[GRAPHIC] [TIFF OMITTED] TR30JY07.001
(iii) for residual strength substantiation, the airplane must be
able to withstand two thirds of the ultimate loads defined in paragraph
(f)(2)(ii) of these special conditions. For pressurized cabins, these
loads must be combined with the normal operating differential pressure.
(iv) If the loads induced by the failure condition have a
significant effect on fatigue or damage tolerance then the effects of
these loads 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[min] and
V[sec] may be based on the speed limitation specified for the remainder
of the flight using the margins defined by Sec. 25.629(b).
Figure 3
Clearance Speed
V[min]=Clearance speed as defined by Sec. 25.629(b)(2).
V[sec]=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[sec].
[GRAPHIC] [TIFF OMITTED] TR30JY07.002
(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 14 CFR part 25 regardless of calculated system
reliability. Where analysis shows the probability of these failure
conditions to be less than 10-9, criteria other than those
specified in this paragraph may be used for structural substantiation
to show continued safe flight and landing.
(g) Failure indications. For system failure detection and
indication, the following apply.
(1) The system must be checked for failure conditions, not
extremely improbable, that degrade the structural capability of the
airplane below the level required by part 25 or significantly reduce
the reliability of the remaining system. As far as reasonably
practicable, the flightcrew must be made aware of these failures before
flight. Certain elements of the control system, such as mechanical and
hydraulic components, may use special periodic inspections, and
electronic components may use daily checks, instead of detection and
indication systems to achieve the objective of this requirement. Such
certification maintenance inspections or daily checks must be limited
to components on which faults are not readily detectable by normal
detection and indication 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 subpart C below 1.25, or flutter margins
below V'', must be signaled to the crew during flight.
(h) Dispatch with known failure conditions. If the airplane is to
be dispatched in a known system failure condition that affects
structural performance, or affects the reliability of the remaining
system to maintain structural performance, then the provisions of these
special conditions must be met, including the provisions of paragraph
(e) for the dispatched condition, and paragraph (f) 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 10-3 per
hour.
[[Page 41433]]
2. Electronic Flight Control System: Control Surface Awareness
In addition to compliance with Sec. Sec. 25.143, 25.671, and
25.672, the following special conditions apply.
(a) The system design must ensure that the flightcrew is made
suitably aware whenever the primary control means nears the limit of
control authority. This indication should direct the pilot to take
appropriate action to avoid the unsafe condition in accordance with
appropriate airplane flight manual (AFM) instructions. Depending on the
application, suitable annunciations may include cockpit control
position, annunciator light, or surface position indicators.
Furthermore, this requirement applies at limits of control authority,
not necessarily at limits of any individual surface travel.
(b) Suitability of such a display or alerting must take into
account that some pilot-demanded maneuvers are necessarily associated
with intended full performance, which may require full 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 nuisance factors. A
monitoring system which might compare airplane motion, surface
deflection, and pilot demand could be useful for eliminating nuisance
alerting.
3. 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 these Special Conditions, the following
definition applies. Critical Functions: Functions whose failure would
contribute to or cause a failure condition that would prevent continued
safe flight and landing of the airplane.
4. Limit Engine Torque Loads for Sudden Engine Stoppage
In lieu of Sec. 25.361(b) the Boeing Model 787-8 must comply with
the following special conditions.
(a) For turbine engine installations, the engine mounts, pylons,
and adjacent supporting airframe structure must be designed to
withstand 1g level flight loads acting simultaneously with the maximum
limit torque loads imposed by each of the following:
(1) Sudden engine deceleration due to a malfunction which could
result in a temporary loss of power or thrust.
(2) The maximum acceleration of the engine.
(b) For auxiliary power unit installations, the power unit mounts
and adjacent supporting airframe structure must be designed to
withstand 1g level flight loads acting simultaneously with the maximum
limit torque loads imposed by each of the following:
(1) Sudden auxiliary power unit deceleration due to malfunction or
structural failure.
(2) The maximum acceleration of the power unit.
(c) For engine supporting structure, an ultimate loading condition
must be considered that combines 1g flight loads with the transient
dynamic loads resulting from each of the following:
(1) Loss of any fan, compressor, or turbine blade.
(2) Where applicable to a specific engine design, any other engine
structural failure that results in higher loads.
(d) The ultimate loads developed from the conditions specified in
paragraphs (c)(1) and (c)(2) are to be multiplied by a factor of 1.0
when applied to engine mounts and pylons and multiplied by a factor of
1.25 when applied to adjacent supporting airframe structure.
5. Design Roll Maneuver Requirement
In lieu of compliance to Sec. 25.349(a), the Boeing Model 787-8
must comply with the following special conditions.
The following conditions, speeds, and cockpit roll control motions
(except as the motions may be limited by pilot effort) must be
considered in combination with an airplane load factor of zero and of
two-thirds of the positive maneuvering factor used in design. In
determining the resulting control surface deflections, the torsional
flexibility of the wing must be considered in accordance with Sec.
25.301(b):
(a) Conditions corresponding to steady rolling velocities must be
investigated. In addition, conditions corresponding to maximum angular
acceleration must be investigated for airplanes with engines or other
weight concentrations outboard of the fuselage. For the angular
acceleration conditions, zero rolling velocity may be assumed in the
absence of a rational time history investigation of the maneuver.
(b) At VA, sudden movement of the cockpit roll control
up the limit is assumed. The position of the cockpit roll control must
be maintained until a steady roll rate is achieved and then must be
returned suddenly to the neutral position.
(c) At VC, the cockpit roll control must be moved
suddenly and maintained so as to achieve a roll rate not less than that
obtained in paragraph (b).
(d) At VD, the cockpit roll control must be moved
suddenly and maintained so as to achieve a roll rate not less than one
third of that obtained in paragraph (b).
Issued in Renton, Washington, on July 18, 2007.
Stephen P. Boyd,
Acting Manager, Transport Airplane Directorate, Aircraft Certification
Service.
[FR Doc. 07-3689 Filed 7-27-07; 8:45 am]
BILLING CODE 4910-13-M