[Federal Register Volume 76, Number 51 (Wednesday, March 16, 2011)]
[Proposed Rules]
[Pages 14341-14346]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2011-6073]


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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 25

[Docket No. NM400 Special Conditions No. 25-11-09-SC]


Special Conditions: Boeing Model 747-8/-8F Airplanes, Interaction 
of Systems and Structures

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Notice of proposed special conditions.

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SUMMARY: This notice proposes to amend Special Conditions No. 25-388-SC 
for the Boeing Model 747-8/-8F airplanes. These special conditions were 
previously issued July 29, 2009, and became effective September 10, 
2009. These special conditions are being amended to include additional 
criteria addressing the Outboard Aileron Modal Suppression System. The 
747-8/-8F 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 their 
effects on the structural performance. These proposed special 
conditions contain the additional safety standards that the 
Administrator considers necessary to establish a level of safety 
equivalent to that established by the existing airworthiness standards. 
Additional special conditions will be issued for other novel or unusual 
design features of the 747-8/-8F airplanes.

DATES: Comments must be received on or before April 15, 2011.

ADDRESSES: Comments on this proposal may be mailed in duplicate to: 
Federal Aviation Administration, Transport Airplane Directorate, 
Attention: Rules Docket (ANM-113), Docket No. NM400, 1601 Lind Avenue 
SW., Renton, Washington 98057-3356; or delivered in duplicate to the 
Transport Airplane Directorate at the above address. All comments must 
be marked Docket No. NM400. Comments may be inspected in the Rules 
Docket weekdays, except Federal holidays, between 7:30 a.m. and 4 p.m.

FOR FURTHER INFORMATION CONTACT: Carl Niedermeyer, FAA, Airframe & 
Cabin Safety Branch, ANM-115, Transport Airplane Directorate, Aircraft 
Certification Service, 1601 Lind Avenue SW., Renton, Washington 98057-
3356; telephone (425) 227-2279; e-mail [email protected].

SUPPLEMENTARY INFORMATION:

Comments Invited

    The FAA invites interested persons to participate in this 
rulemaking by submitting written comments, data, or views. The most 
helpful comments reference a specific portion of the proposed special 
conditions, explain the reason for any recommended change, and include 
supporting data. We ask that you send us two copies of written 
comments.
    We will file in the docket all comments we receive as well as a 
report summarizing each substantive public contact with FAA personnel 
concerning these proposed special conditions. The docket is available 
for public inspection before and after the comment closing date. If you 
wish to review the docket in person, go to the address in the ADDRESSES 
section of this notice between 7:30 a.m. and 4 p.m., Monday through 
Friday, except Federal holidays.
    We will consider all comments we receive on or before the closing 
date for comments. We will consider comments filed late if it is 
possible to do so without incurring expense or delay. We may change the 
proposed special conditions based on comments we receive.
    If you want the FAA to acknowledge receipt of your comments on this 
proposal, include with your comments a pre-addressed, stamped postcard 
on which the docket number appears. We will stamp the date on the 
postcard and mail it back to you.

Background

    On November 4, 2005, The Boeing Company, PO Box 3707, Seattle, WA 
98124, applied for an amendment to Type Certificate Number A20WE to 
include the new Model 747-8 passenger airplane and the new Model 747-8F 
freighter airplane. The Model 747-8 and the Model 747-8F are 
derivatives of the 747-400 and the 747-400F, respectively. Both the 
Model 747-8 and the Model 747-8F are four-engine jet transport 
airplanes that will have a maximum takeoff weight of 970,000 pounds and 
new General Electric GEnx -2B67 engines. The Model 747-8 will have two 
flight crew and the capacity to carry 605 passengers. The Model 747-8F 
will have two flight crew and a zero passenger capacity, although 
Boeing has submitted a petition for exemption to allow the carriage of 
supernumeraries.
    These special conditions were originally issued July 29, 2009, and 
published in the Federal Register on August 12, 2009 (74 FR 40479).

Type Certification Basis

    Under the provisions of Title 14, Code of Federal Regulations (14 
CFR) 21.101, Boeing must show that Model 747-8 and 747-8F airplanes 
(hereafter referred as 747-8/-8F) meet the applicable provisions of 
part 25, as amended by Amendments 25-1 through 25-117, except for 
earlier amendments as agreed upon by the FAA. These regulations will be 
incorporated into Type Certificate No. A20WE after type certification 
approval of the 747-8/-8F.
    In addition, the certification basis includes other regulations, 
special conditions and exemptions that are not relevant to these 
proposed special conditions. Type Certificate No. A20WE will be updated 
to include a complete description of the certification basis for these 
model airplanes.
    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 747-8/-8F 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, or should any other model already 
included on the same type certificate be modified to incorporate 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 747-8/-8F must comply with the fuel vent and exhaust 
emission requirements of 14 CFR part 34 and the noise certification 
requirements of 14 CFR part 36.
    Special conditions, as defined in Sec.  11.19, are issued under 
Sec.  11.38, and become part of the type certification basis under 
Sec.  21.101.

Novel or Unusual Design Features

    The Boeing Model 747-8/-8F 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

[[Page 14342]]

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.
    The Boeing 747-8F airplane exhibits an aeroelastic mode of 
oscillation that is self-excited and does not completely damp out after 
an external disturbance. The sustained oscillation (also known as a 
limit cycle oscillation or limit cycle flutter) is caused by an 
unstable aeroelastic mode that is prevented from becoming a divergent 
oscillation due to one or more nonlinearities that exist in the 
airplane.
    While the sustained oscillation is not divergent, the FAA considers 
it to be an aeroelastic instability. Boeing has proposed the addition 
of an Outboard Aileron Modal Suppression (OAMS) system to the fly-by-
wire (FBW) flight control system to reduce, but not eliminate, the 
amplitude of the sustained oscillation and control the aeroelastic 
instability.
    Section 25.629 requires the airplane to be free of any aeroelastic 
instability, including flutter. It also requires the airplane to remain 
flutter free after certain failures. The regulations do not anticipate 
the use of systems that control flutter modes but do not completely 
suppress them. The use of the OAMS system is a novel and unusual design 
feature that the airworthiness standards do not adequately address. The 
FAA believes such systems can be used to ensure that limit cycle (non-
divergent) flutter is kept to safe levels. Therefore, the FAA proposes 
a special condition that addresses this particular sustained 
oscillation characteristic and provides the necessary standards that 
permit the use of such active flutter control systems.

Applicability

    As discussed above, this proposed special condition is applicable 
to Boeing Model 747-8/-8F airplanes. Should Boeing apply at a later 
date for a change to the type certificate to include another model 
incorporating the same novel or unusual design features, this proposed 
special condition would apply to that model as well under the 
provisions of Sec.  21.101.

Conclusion

    This action affects only certain novel or unusual design features 
of the Boeing Model 747-8/-8F airplanes. 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 this proposed Special Condition is as 
follows:

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

The Proposed Special Conditions

    Accordingly, pursuant to the authority delegated to me by the 
Administrator, the Federal Aviation Administration (FAA) proposes the 
following amendment to Special Conditions 25-388-SC as part of the type 
certification basis for the 747-8/-8F airplanes. The standards in 
Section A have been modified to incorporate the reference to Section C 
and remove ``flutter control systems'' from the applicability of this 
special condition. Section B was already adopted in Special Conditions 
25-388-SC and is included for reference. Comments are invited on the 
amended Section A and the proposed text of Section C, Outboard Aileron 
Modal Suppression System.

A. General

    The Boeing Model 747-8/-8F airplanes are equipped with automatic 
control systems that affect the airplane's structural performance, 
either directly or as a result of a failure or malfunction. The 
influence of these systems and their failure conditions must be taken 
into account when showing compliance with the requirements of Subparts 
C and D of part 25. Except as provided in Section C of this special 
condition, 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, 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.
    1. The criteria defined here only address the direct structural 
consequences of the system responses and performances and cannot be 
considered in isolation; however, they should be included in the 
overall safety evaluation of the airplane. These criteria may in some 
instances duplicate standards already established for this evaluation. 
These criteria are only applicable to structural elements whose failure 
could prevent continued safe flight and landing. Specific criteria that 
define acceptable limits on handling characteristics or stability 
requirements when operating in the system degraded or inoperative mode 
are not provided in this special condition.
    2. 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 the capability of the 
airplane to meet other realistic conditions such as alternative gust or 
maneuver descriptions for an airplane equipped with a load alleviation 
system.
    3. The following definitions are applicable to this special 
condition.
    (a) Structural performance: Capability of the airplane to meet the 
structural requirements of part 25.
    (b) Flight limitations: Limitations that can be applied to the 
airplane flight conditions following an in-flight occurrence and that 
are included in the airplane flight manual (AFM) (e.g., speed 
limitations, avoidance of severe weather conditions).
    (c) Operational limitations: Limitations, including flight 
limitations that can be applied to the airplane operating conditions 
before dispatch (e.g., fuel, payload and Master Minimum Equipment List 
(MMEL) limitations).
    (d) Probabilistic terms: The probabilistic terms (probable, 
improbable, extremely improbable) used in this special condition are 
the same as those used in Sec.  25.1309.
    (e) Failure condition: The term failure condition is the same as 
that used in Sec.  25.1309, however this special condition applies only 
to system failure conditions that affect the structural performance of 
the airplane (e.g., system failure conditions that induce loads, change 
the response of the airplane to inputs such as gusts or pilot actions, 
or lower flutter margins). The system failure condition includes 
consequential or cascading effects resulting from the first failure.

B. Effects of Systems on Structures

    1. General. The following criteria will be used in determining the 
influence of a system and its failure conditions on the airplane 
structural elements.
    2. System fully operative. With the system fully operative, the 
following apply:
    (a) Limit loads must be derived in all normal operating 
configurations of the system from all the limit conditions specified in 
subpart C (or used in lieu of those specified in subpart C), taking 
into account any special behavior of

[[Page 14343]]

such a system or associated functions or any effect on the structural 
performance of the airplane that may occur up to the limit loads. In 
particular, any significant nonlinearity (rate of displacement of 
control surface, thresholds or any other system nonlinearities) must be 
accounted for in a realistic or conservative way when deriving limit 
loads from limit conditions.
    (b) The airplane must meet the strength requirements of part 25 
(i.e., static strength, residual strength), using the specified factors 
to derive ultimate loads from the limit loads defined above. The effect 
of nonlinearities must be investigated beyond limit conditions to 
ensure the behavior of the system presents no anomaly compared to the 
behavior below limit conditions. However, conditions beyond limit 
conditions need not be considered when it can be shown that the 
airplane has design features that will not allow it to exceed those 
limit conditions.
    (c) The airplane must meet the aeroelastic stability requirements 
of Sec.  25.629.
    3. System in the failure condition. For any system failure 
condition not shown to be extremely improbable, the following apply:
    (a) At the time of occurrence, starting from 1-g level flight 
conditions, a realistic scenario including pilot corrective actions, 
must be established to determine the loads occurring at the time of 
failure and immediately after failure.
    (1) For static strength substantiation, these loads multiplied by 
an appropriate factor of safety that is related to the probability of 
occurrence of the failure are ultimate loads to be considered for 
design. The factor of safety (F.S.) is defined in Figure 1.
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[GRAPHIC] [TIFF OMITTED] TP16MR11.003

    (2) For residual strength substantiation, the airplane must be able 
to withstand two thirds of the ultimate loads defined in subparagraph 
3(a)(1). For pressurized cabins, these loads must be combined with the 
normal operating differential pressure.
    (3) 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 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.
    (4) Failures of the system that result in forced structural 
vibrations (oscillatory failures) must not produce loads that could 
result in detrimental deformation of the affected structural elements.
    (b) For continuation of flight, for an airplane in the system 
failed state and considering any appropriate reconfiguration and flight 
limitations, the following apply:
    (1) The 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:
    (i) the limit symmetrical maneuvering conditions specified in Sec.  
25.331 and in Sec.  25.345.
    (ii) the limit gust and turbulence conditions specified in Sec.  
25.341 and in Sec.  25.345.
    (iii) the limit rolling conditions specified in Sec.  25.349 and 
the limit unsymmetrical conditions specified in Sec. Sec.  25.367 and 
25.427(b) and (c).
    (iv) the limit yaw maneuvering conditions specified in Sec.  
25.351.
    (v) the limit ground loading conditions specified in Sec. Sec.  
25.473, 25.491 and 25.493.
    (2) For static strength substantiation, each part of the structure 
must be able to withstand the loads in paragraph (3)(b)(1) of the 
special condition multiplied by a factor of safety depending on the 
probability of being in

[[Page 14344]]

this failure state. The factor of safety is defined in Figure 2.
[GRAPHIC] [TIFF OMITTED] TP16MR11.004

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.

    (3) For residual strength substantiation, the airplane must be able 
to withstand two thirds of the ultimate loads defined in paragraph 
(3)(b)(1) of the special condition. For pressurized cabins, these loads 
must be combined with the normal operating differential pressure.
    (4) If the loads induced by the failure condition have a 
significant effect on fatigue or damage tolerance then their effects 
must be taken into account.
    (5) 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).

[[Page 14345]]

[GRAPHIC] [TIFF OMITTED] TP16MR11.005

BILLING CODE 4910-13-C
    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''.

    (6) 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).
    (c) Consideration of certain failure conditions may be required by 
other sections of part 25 regardless of calculated system reliability. 
Where analysis shows the probability of these failure conditions to be 
less than 10-9, criteria other than those specified in this 
paragraph may be used for structural substantiation to show continued 
safe flight and landing.
    4. Failure indications. For system failure detection and 
indication, the following apply:
    (a) The system must be checked for failure conditions, not 
extremely improbable, that degrade the structural capability below the 
level required by part 25 or significantly reduce the reliability of 
the remaining system. As far as reasonably practicable, the flight crew 
must be made aware of these failures before flight. Certain elements of 
the control system, such as mechanical and hydraulic components, may 
use special periodic inspections, and electronic components may use 
daily checks, in lieu of detection and indication systems to achieve 
the objective of this requirement. These Certification Maintenance 
Requirements (CMRs) must be limited to components that are not readily 
detectable by normal detection and indication systems and where service 
history shows that inspections will provide an adequate level of 
safety.
    (b) 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 flight crew. 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.
    5. 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 2 for the 
dispatched condition, and paragraph 3 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.

C. Outboard Aileron Modal Suppression System

    1. In general, this special condition applies to fly-by-wire active 
flutter suppression systems that are intended to operate on a certain 
type of aeroelastic instability. This type of instability is 
characterized by a low frequency, self-excited, sustained

[[Page 14346]]

oscillation of an aeroelastic vibration mode that is shown to be a 
stable limit cycle oscillation (LCO), with the system operative and 
inoperative. (An LCO is considered ``stable'' if it maintains the same 
frequency and amplitude for a given excitation input and flight 
condition.) In addition, the type of sustained oscillation covered by 
this special condition must not be a hazard to the airplane nor its 
occupants with the active system failed. These systems must be shown to 
reduce the amplitude of the sustained oscillation to acceptable levels 
and effectively control the aeroelastic instability.
    Specifically, the following criteria address the existence of such 
a sustained oscillation on the Boeing Model 747-8/-8F airplanes and the 
Outboard Aileron Modal Suppression (OAMS) system that will be used to 
control it.
    2. In lieu of the requirements contained in Sec.  25.629, the 
existence of a sustained, or limit cycle, oscillation that is 
controlled by an active flight control system is acceptable, provided 
that the following requirements are met:
    (a) OAMS System Inoperative
    (1) The sustained, or limit cycle, oscillation must be shown by 
test and analysis to be stable throughout the nominal aeroelastic 
stability envelope specified in Sec.  25.629(b)(1) with the OAMS system 
inoperative. This should include the consideration of disturbances 
above the sustained amplitude of oscillation
    (b) Nominal Conditions:
    (1) With the OAMS system operative it must be shown that the 
airplane remains safe, stable, and controllable throughout the nominal 
aeroelastic stability envelope specified in Sec.  25.629(b)(1) by 
providing adequate suppression of the aeroelastic modes being 
controlled. All applicable airworthiness and environmental requirements 
should continue to be complied with. Additionally, loads imposed on the 
airplane due to any amplitude of oscillation must be shown to have a 
negligible impact on structure and systems, including wear, fatigue and 
damage tolerance. The OAMS system must function properly in all 
environments that may be encountered.
    (2) The applicant must establish by test and analysis that the OAMS 
system can be relied upon to control and limit the sustained amplitude 
of the oscillation to acceptable levels (per Sec.  25.251) and control 
the stability of the aeroelastic mode. This should include the 
consideration of disturbances above the sustained amplitude of 
oscillation; maneuvering flight, icing conditions; manufacturing 
variations; Master Minimum Equipment List (MMEL) items; spare engine 
carriage; engine removed or inoperative ferry flights; and wear, 
repairs, and modifications throughout the service life of the airplane 
by:
    (i) Analysis to the nominal aeroelastic stability envelope 
specified in Sec.  25.629(b)(1), and
    (ii) Flight flutter test to the VDF/MDF 
boundary. These tests must demonstrate that the airplane has a proper 
margin of damping for disturbances above the sustained amplitude of 
oscillation at all speeds up to VDF/MDF, and that 
there is no large and rapid reduction in damping as VDF/
MDF is approached.
    (iii) The structural modes must have adequate stability margins for 
any OAMS flight control system feedback loop at speeds up to the fail-
safe aeroelastic stability envelope specified in Sec.  25.629(b)(2).
    (c) Failures, Malfunctions, and Adverse Conditions:
    (1) For the OAMS system operative and failed, for any failure, or 
combination of failures not shown to be extremely improbable, and 
addressed by Sec. Sec.  25.629(d), 25.571, 25.631, 25.671, 25.672, 
25.901(c) or 25.1309 that results in LCO, it must be established by 
test or analysis up to the aeroelastic stability envelope specified in 
Sec.  25.629(b)(2) that the LCO:
    (i) is stable and decays to an acceptable limited amplitude once an 
external perturbing force is removed;
    (ii) does not result in loads that would cause static, dynamic, or 
fatigue failure of structure during the expected exposure period;
    (iii) does not result in repeated loads that would cause an 
additional failure due to wear during the expected exposure period that 
precludes safe flight and landing;
    (iv) has, if necessary, sufficient indication of OAMS failure(s) 
and crew procedures to properly address the failure(s);
    (v) does not result in a vibration condition on the flight deck 
that is severe enough to interfere with control of the airplane, 
ability of the crew to read the flight instruments, perform vital 
functions like reading and accomplishing checklist procedures, or to 
cause excessive fatigue to the crew;
    (vi) does not result in adverse effects on the flight control 
system or on airplane stability, controllability, or handling 
characteristics (including airplane-pilot coupling (APC) per Sec.  
25.143) that would prevent safe flight and landing; and
    (vii) does not interfere with the flight crew's ability to 
correctly distinguish vibration from buffeting associated with the 
recognition of stalls or high speed buffet.
    (2) The applicant must show that particular risks such as engine 
failure, uncontained engine, or APU rotor burst, or other failures not 
shown to be extremely improbable, will not adversely or significantly 
change the aeroelastic stability characteristics of the airplane.
    (3) No MMEL dispatch is allowed with the OAMS system inoperative.

    Issued in Renton, Washington on March 9, 2011.
Ali Bahrami,
Manager, Transport Airplane Directorate, Aircraft Certification 
Service.
[FR Doc. 2011-6073 Filed 3-15-11; 8:45 am]
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