[Federal Register Volume 76, Number 130 (Thursday, July 7, 2011)]
[Rules and Regulations]
[Pages 39763-39769]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2011-16295]





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  Federal Register / Vol. 76, No. 130 / Thursday, July 7, 2011 / Rules 

and Regulations  



[[Page 39763]]







DEPARTMENT OF TRANSPORTATION



Federal Aviation Administration



14 CFR Part 25



[Docket No. NM362; Special Conditions No. 25-354A-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: Amended special conditions



-----------------------------------------------------------------------



SUMMARY: These amended special conditions are issued to 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 limit engine torque loads for sudden engine stoppage. 

Special Conditions No. 25-354-SC was issued on July 18, 2007, 

addressing, in part, this condition. We have determined that more 

clarification is needed on the limit engine torque loads for sudden 

engine stoppage special conditions, and have therefore added a new 

requirement. This additional requirement has been applied, via special 

conditions, to other programs. Since applicable airworthiness 

regulations, including those contained in Special Conditions No. 25-

354-SC, do not contain adequate or appropriate safety standards for 

this particular design feature, these amended special conditions 

contain the additional safety standards which the Administrator finds 

necessary to establish a level of safety equivalent to that established 

by the existing standards.



DATES: Effective Date: August 8, 2011.



FOR FURTHER INFORMATION CONTACT: Todd Martin, FAA, Airframe and Cabin 

Safety Branch, ANM-115, Transport Airplane Directorate, Aircraft 

Certification Service, 1601 Lind Avenue, SW., Renton, Washington 98057-

3356, telephone (425-227-1178; facsimile (425-227-1320).



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. Special Conditions 

No. 25-354-SC was issued on July 17, 2007, to address interaction of 

systems and structures, electronic flight control system control 

surface awareness, HIRF protection, limit engine torque loads for 

sudden engine stoppage, and design roll maneuver requirements. Since 

then, it was determined more clarification was needed on the limit 

engine torque loads for sudden engine stoppage special conditions.



Discussion



    The limit engine torque loads for sudden engine stoppage special 

conditions, issued as part of Special Conditions No. 25-354-SC, 

distinguishes between the more common, less severe engine failure 

events, and those rare events resulting from structural failures. 

Paragraph (a) defines limit load conditions for the less severe events, 

and paragraph (c) defines the ultimate load conditions for the more 

severe structural failure events.

    Compliance with paragraph (a) includes, by definition, assessment 

of deformation at limit load, as well as assessment of structural 

integrity at ultimate load. However, since paragraph (c) is defined as 

an ultimate load condition, it only requires assessment of structural 

integrity at ultimate load, and does not require assessment of 

deformation.

    New paragraph (e), therefore, is added to the special condition to 

require assessment of deformation for the structural failures defined 

in paragraph (c).



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 ``the 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.

    If the Administrator finds that the applicable airworthiness 

regulations (i.e., 14 CFR part 25) do not contain adequate or 

appropriate safety standards for the Boeing Model 787-8 because of a 

novel or unusual design feature, special conditions are prescribed 

under the provisions of Sec.  21.16.

    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.

    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.



[[Page 39764]]



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. A special condition is 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.

    This special condition requires that the airplane meet the 

structural requirements of subparts C and D of 14 CFR part 25 when the 

airplane systems are fully operative. The special condition also 

requires 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 flightcrew 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 transmitters 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



[[Page 39765]]



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 25.305 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.

    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 a special condition. This 

special condition contains 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 \2\ 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.

---------------------------------------------------------------------------



    \2\ A nonlinearity is a situation where output does not change 

in the same proportion as input.

---------------------------------------------------------------------------



    This special condition differs from current requirements in that it 

requires that the roll maneuver result from defined movements of the 

cockpit roll control as opposed to defined aileron deflections. Also, 

this special condition requires an additional load condition at design 

maneuvering speed (VA), in which the cockpit roll control is 

returned to neutral following the initial roll input.

    This special condition differs from similar special conditions 

applied to previous designs. This special condition is limited to the 

roll axis only, whereas previous special conditions also included pitch 

and yaw axes. A special condition is 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 

condition is 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.



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 Boeing Model 787-8 airplane. It is not a rule of general 

applicability.



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 Amended Special Conditions



0

Accordingly, pursuance to the authority delegated to me by the 

Administrator, the following amended special conditions (which adds 

paragraph (e) to Special Condition No. 4) are issued as part of the 

type certification basis for the Boeing Model 787-8 airplane regarding 

limit engine torque loads for sudden engine stoppage.



 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 this special condition 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 this special condition is 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 this 

special condition.

    (b) Depending on the specific characteristics of the airplane, 

additional studies may be required that go beyond the criteria provided 

in this special condition in order to demonstrate capability of the 

airplane to meet other realistic conditions such as



[[Page 39766]]



alternative gust conditions or maneuvers for an airplane equipped with 

a load alleviation system.

    (c) The following definitions are applicable to this special 

condition.

    (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 this special condition 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 this special condition, however, 

applies only to system failure conditions that affect 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 this 

special condition, there is no specific reliability requirement for 

the annunciation system required in paragraph (g) of the special 

condition.



    (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] TR07JY11.000



    (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



[[Page 39767]]



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 this 

special condition 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.



[GRAPHIC] [TIFF OMITTED] TR07JY11.001



    (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 this special condition. 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' 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).



Figure 3



Clearance Speed



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'.



[GRAPHIC] [TIFF OMITTED] TR07JY11.002



    (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.



[[Page 39768]]



    (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 this 

special condition 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.



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 condition applies.

    (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 condition.

    (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.

    (e) Any permanent deformation that results from the conditions 

specified in paragraph (c) must not prevent continued safe flight and 

landing.



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 condition.

    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 to 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).





[[Page 39769]]





    Issued in Renton, Washington, on June 23, 2011.

Ali Bahrami,

Manager, Transport Airplane Directorate, Aircraft Certification 

Service, ANM-100.

[FR Doc. 2011-16295 Filed 7-6-11; 8:45 am]

BILLING CODE 4910-13-P