[Federal Register Volume 60, Number 25 (Tuesday, February 7, 1995)]
[Rules and Regulations]
[Pages 7112-7115]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 95-2928]
=======================================================================
-----------------------------------------------------------------------
DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 33
[Docket No. 94-ANE-18; Special Conditions No. SC-33-ANE-08]
Special Conditions; General Electric (GE) Aircraft Engines
Model(s) GE90-75B/-85B/-76B Turbofan Engines
AGENCY: Federal Aviation Administration, DOT.
ACTION: Final special conditions.
-----------------------------------------------------------------------
SUMMARY: These special conditions are issued for the General Electric
(GE) Aircraft Engines Model(s) GE90-75B/-85B/-76B turbofan engines.
These special conditions contain the additional safety standards which
the Administrator considers necessary to establish a level of safety
equivalent to that established by the airworthiness standards of part
33 of the Federal Aviation Regulations (FAR).
EFFECTIVE DATE: March 6, 1995.
FOR FURTHER INFORMATION CONTACT:
Tim Mouzakis at (617) 238-7114 or Karen Grant at (617) 238-7133, Engine
and Propeller Standards Staff, ANE-110, Engine and Propeller
Directorate, Aircraft Certification Service, FAA, New England Region,
12 New England Executive Park, Burlington, Massachusetts 01803-5229;
fax (617) 238-7199.
SUPPLEMENTARY INFORMATION:
Background
On December 16, 1991, General Electric Aircraft Engines applied for
type certification of Model(s) GE90-75B/-85B/-76B turbofan engines.
These engines incorporate a first stage fan blade manufactured using
carbon graphite composite material. This unusual design feature results
in the GE90 fan blade having significant differences in material
property characteristics when compared to conventionally designed fan
blades using non-composite materials. For example, the probability that
a composite fan blade will fail below the inner annulus flowpath line
may be highly improbable, questioning the appropriateness of the
requirement contained in Sec. 33.94(a)(1) to show blade containment
after a failure of the blade at the outermost retention feature.
The current requirements of Sec. 33.94 are based on metallic blade
characteristics and service history, and are not appropriate for the
unusual design features of the composite fan blade found on the GE90
series turbofan engines. The FAA has determined that a more realistic
blade out test will be achieved with a fan blade failure at the inner
annulus flowpath line (only the airfoil) instead of the outermost
retention feature as is currently required by Sec. 33.94(a)(1).
The FAA has also determined that the composite fan blades
construction presents other factors that must be considered. Tests and
analyses must account for the effects of in-service deterioration of,
manufacturing and materials variations in, and environmental effects on
the composite material. Further, tests and analyses must show that a
lightning strike on the composite fan blade will not result in a
hazardous condition to the aircraft, and that the engine will meet the
requirements of Sec. 33.75. Therefore, these special conditions are
additional requirements which the Administrator considers necessary to
establish a level of safety equivalent to that established by the
Airworthiness Standards of part 33.
Type Certification Basis
Under the provisions of Sec. 21.101 of the Federal Aviation
Regulations (FAR), General Electric Aircraft Engines must show that the
Model(s) GE90-75B/-85B/-76B turbofan engines meet the requirements of
the applicable regulations in effect on the date of the application.
Those Federal Aviation Regulations are Sec. 21.21, as amended through
Amendment 21-68, August 10, 1990, and part 33, as amended 33-14, August
10, 1990.
The Administrator finds that the applicable airworthiness
regulations in part 33, as amended, do not contain adequate or
appropriate safety standards for the General Electric Aircraft Engines
Model(s) GE90-75B/-85B/-76B turbofan engines because of unique design
criteria. Therefore, the Administrator prescribes special conditions
under the provisions of Sec. 21.16 to establish a level of safety
equivalent to that established in the regulations.
Special conditions, as appropriate, are issued in accordance with
Sec. 11.49 of the FAR after public notice and opportunity for comment,
as required by Secs. 11.28 and 11.29(b), and become part of the type
certification basis in accordance with Sec. 21.101(b)(2).
Discussion of Comments
Interested persons have been afforded the opportunity to
participate in the making of these special conditions. Due
consideration has been given to comments received.
Two commenters express no objection to the adoption of these
special conditions as proposed.
Two commenters cite the apparent departure by the FAA from its
general practice of involving industry prior to effecting significant
changes to certification requirements, and recommend that the FAA
evaluate the proposed changes in harmony with industry through the
Aviation Rulemaking Advisory Committee (ARAC).
The FAA has not determined that these special conditions will form
the basis to a rulemaking change to amend 14 CFR part 33. These special
conditions prescribe for a specific design, the testing and analyses
necessary to achieve an equivalent level of safety. The FAA may
consider whether it is necessary to revise Sec. 33.94 to include the
requirements of these special conditions. The ARAC may be used to
gather industry and public participation in that rulemaking project.
For this specific application for type certification, however, the FAA
has followed the rulemaking procedures provided by 14 CFR part 11 that
allow for industry and public comment.
Two commenters state that applying the maximum load criteria used
for propellers to a fan blade, with significantly different mechanical
arrangement and dynamic behavior, is technically unjustified.
The FAA disagrees. The two times maximum load criteria test is
designed to show the capability of the fan blade retention system to
withstand without separation centrifugal loads significantly greater
than will be seen in service. A safety factor of two is a reasonable
safety factor as demonstrated by its success in propeller applications.
The blade and its retention system must be capable of retaining the
blade under this load condition.
Two commenters state that the additional requirements, in
conjunction with any available analyses, cannot guarantee that the
failure probability will be extremely improbable. Inherent
characteristics of complex composite hardware design, latent defects
and susceptibility to manufacturing variations, and nonconformance are
identified as reasons for the statement.
The FAA agrees in part. The FAA has reviewed its position and
concurs with the commenters that a failure [[Page 7113]] probability of
extremely improbable can not be guaranteed. However, the FAA believes
that the applicant has constructed a test program that demonstrates the
blade retention features have sufficiently improved reliability to
provide an equivalent level of safety to that provided by Sec. 33.94.
While extensive testing is required for material certification in
accordance with Sec. 33.15 to determine material characteristics and
the effects of defects on blade life, additional test requirements were
established within the compliance plan to determine the effect of
defects and manufacturing variations on material capability.
One commenter suggests adding an additional paragraph to these
special conditions as follows:
``(a)(3) By appropriate test and analysis it must be shown that the
most adverse blade vibratory stresses, as determined per Sec. 33.83,
will not result in failure of the fan blade retention system when
consideration is given to the most limiting manufacturing defect which
could go undetected.''
The FAA disagrees with the commenter that the suggested paragraph
be added, as these considerations are well within the interpretation of
Sec. 33.83 and no additional safety standards are deemed necessary.
One commenter suggests adding an additional paragraph to the
special condition to minimize the risk of hazard which would result
from potential failure of the fan blade retention system as follows:
``(a)(4) Although the above test requires release of the fan blade
at the inner flowpath, additional testing and/or analysis shall be
performed to define the engine behavior for the case of a fan blade
release at the outermost retention groove. The data obtained shall be
used when establishing:
(i) Any installation limitations to be included on the Type
Certificate Data Sheet; and,
(ii) Load requirements of Sec. 33.23.''
The FAA disagrees. As stated in Sec. 33.75, Safety Analysis, the
applicant must consider all probable malfunctions which will cause the
engine to catch fire, burst, generate loads greater than those ultimate
loads specified in Sec. 33.23(a), or lose the capability of being shut
down. These special conditions also require such analyses and tests to
show that the failure of the fan blade retention system is not a
probable malfunction. Establishment of the maximum stop-start stress
cycles for the blade retention system is also required to assure the
structural integrity of the blade attachment system.
One commenter states that the requirements should show that the
failure rate of the fan blade retention system, for any cause, during
the service life of the engine, be extremely improbable and can not be
established at the time of type design approval for a new technology
composite.
The FAA agrees in part. While the FAA agrees that a failure
probability of extremely improbable can not be guaranteed, the FAA
remains receptive to advances in technology, approaches, and new test
methods which adequately simulate those effects typically verified by
in-service experience. Further, the FAA believes that these same
principles have been successfully used by engine manufacturers to
ensure the airworthiness of rotor structural parts. It should be
recognized that failure to demonstrate acceptable reliability of the
blade retention features, results in non-compliance with these special
conditions and that would require testing to occur at the outer most
retention groove.
Two commenters suggest the energy levels and trajectories of any
particles that would penetrate the engine cases by conducting an engine
test in accordance with the test conditions of current Secs. 33.94(a)
and 33.94(b) be defined in the Engine Installation Manual or on the
Engine Type Certificate Data Sheet. The definition of results should
also include determination of the loads that would be transmitted
through the engine to airframe interface. One commenter states that the
energy levels, trajectories and loads must be included in each airplane
type's design precautions taken to minimize the hazards in the event of
an engine rotor failure, as required by current FAR 25.903 and JAR
25.903.
The FAA agrees that the requirements for defining energy levels,
trajectories of particles, and a resultant loads already exist in
Secs. 33.19(a) and 33.23. The FAA also agrees that if such energy
levels, trajectories, and resultant loads are defined, the appropriate
data should be included in the Engine Installation Manual. The FAA does
not agree with the commenters suggestion relative to complying with
Secs. 33.94(a) and 33.94(b) in addition to these special conditions.
These special conditions provide safety standards which apply to the
composite blade design as an alternative to the requirements of
Sec. 33.94. The applicant must demonstrate reliability of the blade
root and the blade retention system.
One commenter criticizes the explanations and logic presented for
justification of these proposed special conditions. The commenter cites
that there was insufficient information in the notice by which to test
the validity of the FAA's determination.
The FAA disagrees. The notice of proposed special condition
identifies two bases on which the FAA determined that the current
requirements of part 33 do not provide adequate or appropriate safety
standards because of the novel or unusual design of the GE90 engine.
The FAA also determined that additional safety standards were needed to
ensure that the GE composite fan blades met an equivalent level of
safety established by Sec. 33.94. Given the number and the nature of
the comments received, the FAA believes that the notice gave an
adequate description of the proposed action to allow critical comment
on the basis for that action.
One commenter states that they do not believe that use of graphite
composite material for a turbofan blade retention system warrants a
departure from the current requirements of Sec. 33.94.
The FAA disagrees. The FAA supports the use of composite technology
and the necessary methods of testing and analyses to show that the
product meets an equivalent safety standard as established by
Sec. 33.94.
One commenter states that the demonstration means for showing
``extremely improbable'' should be specifically part of these proposed
special conditions. The commenter suggests to establish and define a
methodology by which to rigorously assess the probability of fan blade
retention system failure as extremely improbable, and by which to
assess the associated level of confidence in the assessment,
particularly at the time of initial certification.
The FAA agrees in part. The FAA agrees that the assessment of the
fan blade retention system should be conducted rigorously, but
disagrees with the need to establish and define a methodology in these
special conditions. The FAA believes it should not define a specific
means to meet a safety standard, or publish an applicant's proprietary
methodology. To publish a specific demonstration means would presume
the FAA has predetermined the composite blade material property
characterization. The methodology for assessing the fan blade retention
system will be proposed by the applicant, and will be evaluated by the
FAA.
One commenter states that lightning test conditions should be
specifically identified in the special condition.
The FAA disagrees. Existing regulatory guidance material and
standard industry practices for lightning [[Page 7114]] tests may be
used to develop appropriate test criteria.
One commenter suggests that the term ``inner annulus flowpath
line'' be substituted for ``inner flowpath diameter'' to eliminate
ambiguity of definition.
The FAA concurs. The inner annulus flowpath line provides a better
description of the flowpath contour because flowpath diameter suggests
a line of constant radius. These Final Special Conditions will be
revised to include this term.
One commenter states it is an issue of unnecessary additional risk
that, in the absence of full compliance to Sec. 33.94, these proposed
special conditions are insufficient in scope and detrimental to
aviation safety.
The FAA disagrees. The FAA has concluded that upon compliance with
all of the requirements of these special conditions, together with
additional testing beyond that typically employed for metallic blades
within the scope of 14 CFR part 33, an equivalence to the safety
standard provided in Sec. 33.94 has been achieved and no additional
risk has been assumed.
One commenter states that the most significant feature of the
notice is the proposed probability of fan blade retention system
failure of ``extremely improbable'' is a reduction in severity of the
effects of a blade failure.
The FAA agrees. The FAA recognizes that certain loads associated
with a blade release at the inner annulus flowpath line may be less
than the loads associated with release of a fan blade at the outermost
retention. Those loads imparted to the engine mount system based on the
inner annulus flowpath line will be identified in the Engine
Installation Manual. Since there is potential for a reduction in
certain loads, it is imperative that the blade retention system
demonstrates sufficiently improved reliability to provide an equivalent
level of safety to that provided by Sec. 33.94.
One commenter requested on what basis has it been decided that a
failure along the inner flowpath line is the most critical for failures
which are not assessed as being extremely improbable.
The FAA selected the inner annulus flowpath line as the critical
location for blade release based on design, blade stresses, and
demonstrated fatigue and impact testing.
One commenter states that these proposed special conditions make no
mention of the design and construction requirements of either
Sec. 33.19 relating to containment design and uncontained blade
fragments, or Sec. 33.23 relating to mounting attachments and
structure.
The FAA concluded that the requirements of Secs. 33.19 and 33.23
were adequate and appropriate when applied to this design of the GE90
engine, and no additional special conditions were necessary.
One commenter suggests that these special conditions should also
address the effects of possible detachment of those metallic portions
of the blade.
The FAA disagrees. These special conditions provide an alternative
to the release failure location on the blade. The metal to composite
blade bonding capability has been addressed through tests conducted
under 14 CFR part 33. There were no additional special conditions that
are required.
One commenter suggests that the text of these proposed special
conditions paragraph (a), has been mis-compiled.
The FAA concurs. The intent of the paragraph (a) is to identify the
location of the release point for the fan blade containment test and to
prescribe the additional safety standards to be demonstrated. These
special conditions will be modified by reorganizing paragraph (a) to
more clearly express this intent.
One commenter states that some re-wording is also necessary to make
it clear that the fan blade test must be conducted as a full engine
test.
The FAA concurs. These special conditions will be modified to
incorporate this change.
One commenter states that these special conditions ought to make
more visible how there can be meaningful confidence in ``extremely
improbable'' as the assessed probability of fan blade retention system
failure if the stress levels are not so conservative as to result in an
infinite fatigue life.
The FAA disagrees. The intent is to assure that within the service
life of the blade, that the fan blade retention system is not likely to
fail due to manufacturing and material variations, in-service
deterioration, and environmental effects.
One commenter asks how will it be established that any large bird
ingestion is not a possible cause of fan blade retention system
failure, a mode of failure that is likely to be much more severe than
an airfoil only fan blade containment tests.
The damage effects on the blade retention system will be
substantiated by developmental and certification testing. It is
incumbent upon the applicant to demonstrate that the blade attachment
system is designed to withstand the affects of an eight pound bird
impact on the blade airfoil, and is less severe than the effects from
fan blade release.
One commenter requests a definition of ``without failure,'' with
regard to the two times centrifugal load test.
The FAA definition for ``without failure'' in this context is to
demonstrate the blade root is retained within the disk dovetail slot,
and that there are no conditions present which would indicate impending
release.
One commenter suggests relative to paragraph (a)(2) of the proposed
special conditions, that there is a need for explicit reference to
consideration of both high cycle and low cycle fatigue during start
stop stress cycles.
The FAA concurs. The determination of the life cycle of the
composite fan blade must include the effects of combined high cycle and
low cycle fatigue with enhanced load factors. These special conditions
will be modified to include the requirement for high cycle and low
cycle fatigue tests.
One commenter requests clarification of the term ``extremely
improbable.''
For the purpose of these special conditions, ``extremely
improbable'' refers to the unlikelihood that a failure will occur
during the engine's operational life.
One commenter questions why paragraph (d) of these proposed special
conditions is applicable only to the tests and analyses required by
paragraphs (a)(1) and (a)(2) of the proposed special conditions.
The effects of in-service deterioration, manufacturing and material
variations, and environmental effects must be accounted for during the
centrifugal load test and in lifting determinations. The intent is to
determine the effects on material capability under centrifugal loads
significantly greater than will be seen in service. Combined high cycle
and low cycle tests will further determine the effects on material
capability. The blade releases demonstration, however, may or may not
be conducted accounting for these effects.
After careful review of the available data, including the comments
noted above, the FAA determined that air safety and the public interest
require the adoption of these special conditions as proposed with the
changes as noted above.
Conclusion
This action affects only General Electric Aircraft Engines on
Model(s) GE90-75B/-85B/-76B turbofan engines. It is not a rule of
general applicability and affects only the manufacturer who applied to
the FAA for approval of these engines containing this novel or unusual
design feature. [[Page 7115]]
List of Subjects in 14 CFR Part 33
Air transportation, Aircraft, Aviation safety, Safety.
The authority citation for these special conditions continues to
read as follows:
Authority: 49 U.S.C. App. 1354(a), 1421, 1423; 49 U.S.C. 106(g);
and 14 CFR 11.49 and 21.16.
The Special Conditions
Accordingly, pursuant to the authority delegated to me by the
Administrator, the following special conditions are issued as part of
the type certification basis for the General Electric Aircraft Engines
Model(s) GE-90-75B/-85B/-76B turbofan engines:
(a) In lieu of the fan blade containment test with the fan blade
failing at the point specified in Sec. 33.94(a)(1), conduct the
following:
(1) An engine fan blade containment test with the fan blade
failing at the inner annulus flowpath line.
(2) The following must be shown by test and analyses, or other
methods acceptable to the Administrator, that:
(i) The disk and fan blade retention system can withstand
without failure a centrifugal load equal to two times the maximum
load which the engine could experience within approved operating
limitations, and
(ii) By a procedure approved by the Administrator, an operating
limitation must be established which specifies the maximum allowable
number of start-stop stress cycles for the fan blade retention
system. The stress cycle shall include the combined effects of high
cycle and low cycle fatigue. The fan blade retention system includes
the portion of the fan blade from the inner annulus flowpath line
inward to the blade dovetail, the blade retention components and the
fan disk and fan blade attachment features.
(b) It must be shown that the probability of fan blade retention
system failure, for any cause, during the service life of the engine
to be extremely improbable.
(c) It must be shown by test or analysis that a lightning strike
to the composite fan blade structure will not result in a hazardous
condition, and that the engine will meet the requirements of
Sec. 33.75.
(d) The tests and analyses required by (a)(2)(i) and (a)(2)(ii)
of these special conditions must account for the effects of in-
service deterioration, manufacturing and material variations, and
environmental effects.
Issued in Burlington, Massachusetts, on February 1, 1995.
James C. Jones,
Acting Manager, Engine and Propeller Directorate, Aircraft
Certification Service.
[FR Doc. 95-2928 Filed 2-2-95; 9:32 am]
BILLING CODE 4910-13-M
