[Federal Register Volume 72, Number 145 (Monday, July 30, 2007)]
[Rules and Regulations]
[Pages 41428-41433]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 07-3689]


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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 (V<INF>A</INF>), 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 (V<INF>c</INF>/M<INF>c</INF>), 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 V<INF>c</INF>/M<INF>c</INF>, 
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<SUP>-3</SUP> 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<SUP>-</SUP>\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<SUP>'</SUP> 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<SUP>-9</SUP>, 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<SUP>''</SUP>, 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<SUP>-3</SUP> 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 V<INF>A</INF>, 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 V<INF>C</INF>, 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 V<INF>D</INF>, 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