CFR NPRM

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[Federal Register: August 15, 1977 (Volume 42, Number 157)]
[Page 41236]

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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 25
[Docket No. 16280; Notice No. 77-15]

Transport Category Airplane Fatigue Regulatory Review Program; Fatigue Proposals

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AGENCY: Federal Aviation Administration, DOT
ACTION: Notice of Proposed Rulemaking

SUMMARY: This notice proposes to amend the standards applicable to the type certification of transport category airplanes by revising the structural fatigue evaluation requirements and by incorporating related compliance provisions. Adoption of these proposals would improve and update the current standards by taking into account state-of-the-art developments and accumulated service experience. In general, the notice is based on proposals discussed at the Transport Category Airplane Fatigue Regulatory Review Conference during March 15-17, 1977.
DATES: Comments must be received on or before November 14, 1977.
ADDRESSES: Send comments on the proposals in duplicate to: Federal Aviation Administration, Office of the Chief Counsel, Attn.: Rules Docket (AGC-24), Docket No. 16280, 800 Independence Avenue, SW., Washington, D.C. 20591.
FOR FURTHER INFORMATION CONTACT: Adolfo O. Astorga, Airworthiness Review Branch (AFB-910), Flight Standards Service, Federal Aviation Administration, 800 Independence Avenue SW., Washington, D.C. 20591; Telephone 202-755-8714.
SUPPLEMENTARY INFORMATION:

Comments Invited
Interested persons are invited to participate in the making of the proposed rule by submitting such written data, views, or arguments as they may desire. Communications should identify the regulatory docket or notice number and be submitted in duplicate to the address specified above. All communications received on or before November 14, 1977, will be considered by the Administrator before taking action on the proposed rule. The proposals contained in this notice may be changed in the light of comments received. All comments submitted will be available, both before and after the closing date for comments, in the Rules Docket for examination by interested persons. A report summarizing each FAA-public contact concerned with the substance of this proposal will be filed in the Rules Docket.

Availability of Additional Copies of Notice
Any person may obtain a copy of this notice of proposed rulemaking (NPRM) by submitting a request to: Federal Aviation Administration, Office of Public Affairs, Attn: Public Information Center (APA-430), 800 Independence Avenue SW., Washington, D.C. 20591; Telephone 202-426-8058. Each communication must identify the notice number of this NPRM. Persons interested in being placed on a mailing list for future NPRM's should also request a copy of Advisory Circular No. 11-2 which describes the application procedure.

Background and Discussion of Proposals
During recent years, there have been significant state-of-the-art and industry-practice developments in the area of structural fatigue and fail-safe-strength evaluation of transport category airplanes. Recognizing that these developments could warrant some revision of existing fatigue requirement contained in Sections 25.571 and 25.573 of Part 25 of the Federal Aviation Regulations, the FAA, on November 18, 1976, gave notice of its Transport Category Airplane Fatigue Regulatory Review Program and invited interested persons to submit proposals to amend those requirements (see 41 FR 50956). Subsequently, the FAA convened a Transport Category Airplane Fatigue Regulatory Review Conference during March 15-17, 1977, in Arlington, VA., to obtain the views of all concerned on the proposals submitted for the review.
Participants in the Review Conference discussed the proposals submitted for the review. These proposals and the related discussions form the basis for the FAA's belief that a comprehensive revision of the structural fatigue evaluation standards of Sections 25.571 and 27.573 is warranted. To that end, it is proposed to substantially revise Section 25.571, to delete Section 25.573 (but expand the scope of Section 25.571 to cover the substance of the deleted section), and to add a new Appendix H to Part 25 containing compliance provisions related to proposed Section 25.571. A discussion of significant aspects of the proposed changes follows:
1. Requirement for damage-tolerance (fail-safe) fatigue evaluation. Under current Sections 25.571 and 25.573, applicants for a type certificate must show that critical parts of the airplane structure have either adequate fatigue strength (capability to withstand the repeated loads of variable magnitude expected in service) or adequate fail-safe strength (capability to withstand specified loads after fatigue failure of a principal structural element). Service experience has indicated that structure designed to these criteria are nevertheless subject to unpredictable deterioration in service. Fatigue cracks may develop despite exhaustive fatigue analysis and testing; corrosion may weaken the structure; and accidental damage may occur.
Manufacturers of modern transport category airplanes have successfully dealt with this problem by using, wherever practical, the approach known in the industry as "damage-tolerance (fail-safe)" design. The objective of this approach is to design the structure in such a manner that, should damage occur, the structure would retain adequate residual strength until the damage is detected (and repaired) in the course of a manufacturer-developed inspection program. The inspection program is based on analytical predictions, testing of complete structure or subcomponents, and previous operational and design experience. If the damage-tolerance (fail-safe) approach is found to be impractical for particular structure, a conventional fatigue (safe-life) design is used.
The FAA believes that the use, where practical, of structural design based on a damage-tolerance (fail-safe) evaluation would raise the current level of structural reliability and would further reduce the existing extremely low probability of catastrophic failure.
Accordingly, proposed Section 25.571(a), in part, would require the use, where practical, of a damage-tolerance (fail-safe) evaluation of critical structure; proposed Section 25.571(b) would set forth the basic criteria for that evaluation, including the applicable loading conditions; and proposed Section 25.571(c) would allow the use of a fatigue (safe-life) evaluation if the applicant establishes that a damage-tolerance (fail-safe) evaluation under proposed Section 25.571(b) would be impractical for specific structure. In addition, proposed Sections 25.571(d) and 25.629(d)(4)(v) would be revised to refer to the loading conditions in proposed Section 25.571(b) but would otherwise remain substantively unchanged.
Proposed Section 25.571(b) would require, among other things, that multiple site damage be considered during damage-tolerance (fail-safe) evaluation of critical structure. Service experience has shown that when cracks occur, they can propagate in more than one location simultaneously.
It should be noted that the loading conditions set forth in proposed Section 25.571(b) are expanded relative to those in current Section 25.571(c). The FAA believes that the proposed loading conditions are more representative of actual in-service airplane usage.
Moreover, based on a review of the probability of attaining limit load during actual service, the FAA believes that it is necessary to specify a residual static strength level of 100 percent of limit load in place of the 80 percent of limit load in place of the 80 percent of limit load specified in current Section 25.571(c) and the 1.15 dynamic effects factor specified in current Section 25.571(e). This proposed revision is covered in the loading conditions of proposed Section 25.571(b).
In addition, proposed Section 25.571(b)(5)(ii) would eliminate the 1.33 factor applied to the pressurized cabin load condition in current Section 25.571(e)(2) in favor of a loading condition based on the expected external aerodynamic pressure in 1g flight combined with a cabin differential pressure equal to 1.1 times the normal operating pressure without any other load. The FAA believes that this is a more realistic loading condition for damage-tolerance (fail-safe) evaluation since the inspection procedures the manufacturer would be required to establish under proposed Section 25.571(a)(3) are keyed to a specified number of cycles at normal operating differential pressure.
2. New requirement for damage-tolerance (discrete source) evaluation. The FAA has determined, on the basis of service experience, that there is a need to design transport category airplanes so that the structure can withstand likely structural damage due to the impact of birds or the impact of fragments from failed propellers, engines, or high-energy rotating machinery. Accordingly, proposed Section 25.571(e) would require that following such an impact the airplane be capable of successfully completing the flight.
3. Deletion of current Section 25.573. The FAA believes that there is no longer a need to distinguish between landing gear and other structure with respect to damage-tolerance or fatigue evaluation. Therefore, as proposed, Section 25.573 would be deleted, and Section 25.571 would apply to all parts of the structure which could contribute to catastrophic failure of the airplane. To avoid possible confusion, the term "landing gear" would be added to the list of parts in proposed Section 25.571(a) that would be specifically identified as covered under Section 25.571. It should be noted that the term "empennage" would also be added to the list to make it clear that the empennage warrants the same evaluation as the other parts listed.
4. New Appendix H to Part 25. To provide detailed guidance for showing compliance with the general requirements set forth in proposed Section 25.571, the FAA proposes to add a new Appendix H to Part 25. The compliance provisions in this Appendix would supplement the provisions in proposed Section 25.571 dealing with damage-tolerance (fail-safe), fatigue (safe-life), and damage-tolerance (discrete source) evaluations of structure.
Under Section H25.2(a) of proposed Appendix H, the type certificate applicant would be allowed to use probabilistic methods of damage-tolerance evaluation (such as risk analysis) rather than the more commonly used deterministic methods if the applicant shows that the loss of the airplane would be extremely improbable and that the statistical data employed in the analysis is based on operational experience or tests of similar structure. The FAA believes that in certain instances the use of probabilistic methods would enable a more realistic assessment of damage-tolerance, without compromising safety.
Also under Section H25.2(a) of proposed Appendix H, the applicant would be allowed to apply the damage-tolerance (fail-safe) evaluation approach to both single load path and multiple load path structure. The FAA believes that the applicant can, by sufficient analysis and testing, establish that a single load path structure has sufficiently slow crack growth properties so that, if a crack were to develop, it would be discovered during a properly designed inspection program.
5. Sonic fatigue evaluation for all turbine-engine-powered airplanes. Under current Section 25.571(a) sonic fatigue evaluation of the structure is prescribed for turbojet powered airplanes only. The FAA believes that other turbine-engine-powered airplanes are also subject to sonic fatigue with essentially the same effect on safety. Accordingly, proposed Section 25.571(a) would require sonic fatigue evaluation under proposed Section 25.571(d) for all turbine-engine-powered air planes.
6. Establishment of inspection or other procedures in accordance with Section 25.1529. The FAA believes that a vital and necessary element of both the damage-tolerance evaluation approach and the fatigue evaluation approach is an inspection program aimed at detecting (and permitting repair of) initial damage in sufficient time to prevent catastrophic failure. Accordingly, proposed Section 25.571(a)(3) would require the manufacturer to establish inspections or other procedures, based on the evaluation approach selected, as necessary to prevent catastrophic failure.
7. Consideration of temperatures and humidities expected in service. The FAA believes that variations in both temperature and humidity in the airplane's operating environment may have a significant effect on the damage-tolerance and fatigue characteristics of conventional metallic materials, and of composite materials, used in critical structures. Accordingly, proposed Section 25.571(a)(1)(i) would require consideration of these effects in the required evaluations.

Drafting Information
The principal authors of this document are Irving Fagin, Flight Standards Service, and Samuel Podberesky, Office of the Chief Counsel.

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The Proposed Amendment
Accordingly, the Federal Aviation Administration proposes to amend Part 25 of the Federal Aviation Regulations (14 CFR Part 25) as follows:
1. By Revising Section 25.571 to read as follows:

Section 25.571 Damage-tolerance and fatigue evaluation of structure
(a) General. The strength, detail design, and fabrication of the airframe structure must be such as to avoid catastrophic failure due to fatigue, corrosion, or accidental damage throughout its operational life. This must be shown by an evaluation conducted in accordance with the provisions of paragraphs (b) and (c) of this section, except as specified in paragraph (c) of this section, for each part of the structure which could contribute to such failure (including wing, empennage, control surfaces and their systems, the fuselage, landing gear, and their related primary attachments). For turbine-engine-powered airplanes, those parts must also be evaluated under paragraph (d) of this section. Appendix H of this part contains compliance provisions related to the requirements of this section. In addition, the following apply:
(1) Each evaluation required by this section must include--
(i) The typical loading spectra, temperatures, and humidities expected in service;
(ii) The identification of principal structural elements and detail design points, the failure of which could cause catastrophic failure of the airplane; and
(iii) An analysis, supported by test evidence, of the principal structural elements and detail design points identified in paragraph (a)(1)(ii) of this section.
(2) The service history of airplanes of similar structural design, taking due account of differences in operating conditions and procedures, may be used in the evaluations required by this section.
(3) Based on the evaluations required by this section, inspections or other procedures must be established as necessary to prevent catastrophic failure, and must be included in the maintenance manual required by Section 25.1529.
(b) Damage-tolerance (fail-safe) evaluation. The evaluation must include a determination of the probable locations and modes of damage due to fatigue, corrosions, or accidental damage. The determination must be by analysis supported by test evidence and service experience. Damage at multiple sites due to prior fatigue exposure must be included where the design is such that this type of damage can be expected to occur. The evaluation must incorporate repeated load and static analyses supported by test evidence. The extent of damage for residual strength evaluation at any time within the operational life must be consistent with the initial detectability and subsequent growth under repeated loads. The residual strength evaluation must show that the remaining structure is able to withstand loads (considered as static ultimate loads) corresponding to the following conditions:
(1) The limit symmetrical maneuvering conditions specified in Section 25.337 at VC and in Section 25.345.
(2) The limit gust conditions specified in Sections 25.341 and 25.351(b) at the specified speeds up to VC, and in Section 25.345.
(3) The limit rolling conditions specified in Section 25.349 and the limit unsymmetrical conditions specified in Sections 25.367 and 25.427, at speeds up to VC.
(4) The limit yaw maneuvering conditions specified in Section 25.351(a) at the specified speeds up to VC.
(5) For pressurized cabins, the following conditions:
(i) The normal operating pressure combined with the expected external aerodynamic pressures must be applied simultaneously with the flight loading conditions specified in paragraphs (b)(1) through (4) of this section if they have a significant effect.
(ii) The expected external aerodynamic pressures in 1g flight combined with a cabin differential pressure equal to 1.1 times the normal operating pressure without any other load.
(6) For landing gear and directly-affected airframe structure, the limit ground loading conditions specified in Sections 25.473, 25.491, and 25.493.
If significant changes in structural stiffness or geometry, or both, follow from a structural failure, or partial failure, the effect on damage tolerance must be further investigated.
(c) Fatigue (safe-fail) evaluation. Compliance with the damage-tolerance requirements of paragraph (b) of this section is not required if the applicant establishes that their application for particular structure is impractical. This structure must be shown by analysis, supported by test evidence, to be able to withstand the repeated loads of variable magnitude expected during its service life. Any cracking must be detectable by standard inspection methods. Appropriate safe-life scatter factors must be applied.
(d) Sonic fatigue strength. It must be shown by analysis, supported by test evidence, or by the service history of airplanes of similar structural design and sonic excitation environment, that:
(1) Sonic fatigue cracks are not probable in any part of the flight structure subject to sonic excitation; or
(2) Catastrophic failure, caused by sonic cracks is not probable, assuming that the loads prescribed in paragraph (b) of this section are applied to all areas affected by those cracks.
(e) Damage-tolerance (discrete source) evaluation. The airplane must be capable of successfully completing a flight during which likely structural damage occurs as a result of:
(1) Impact with a 4-pound bird at likely operational speeds at altitudes up to 8,000 feet;
(2) Propeller and uncontained fan blade impact;
(3) Uncontained engine failure; or
(4) Uncontained high energy rotating machinery failure.
The damaged structure must be able to withstand the static loads (considered as ultimate loads) which are reasonably expected to occur on the flight. Corrective action to be taken by the pilot following the incident, such as limiting maneuvers, avoiding turbulence, and reducing speed must be considered. Dynamic effects on static loads must be considered. Dynamic effects on static loads must be considered. However, if significant changes in structural stiffness or geometry, or both, follow from a structural failure, or partial failure, the effect on damage tolerance must be further investigated.

Section 25.573 [Deleted]
2. By deleting Section 25.573 and marking it "[Reserved]".

Section 25.629 [Amended]
3. By amending Section 25.629(d)(4)(v) by deleting "Section 25.571(c)" and inserting in its place "Section 25.571(b)".
4. By adding a new Appendix H to read as follows:

Appendix H -- Damage-Tolerance and Fatigue Evaluation of Structure
H25.1 Introduction. The provisions contained in this Appendix are considered by the FAA in determining compliance with the damage-tolerance and fatigue requirements of Section 25.573. Although a uniform approach to the evaluation required by Section 15.571 is desirable, it is recognized that in such a complex field new design features and methods of fabrication, new approaches to the evaluations, and new configurations may necessitate variations and deviations from the procedures described in this Appendix.
Damage-tolerance design is required, unless it entails such complications that an effective damage-tolerant structure cannot be achieved within the limitations of geometry or inspectability. Under these circumstances, a design that complies with the fatigue evaluation (safe-life) requirements must be used. Typical examples of structure that may not be conducive to damage-tolerance design are landing gear and engine mounting attachments.
Experience with the application of methods of fatigue evaluation indicates that a test background must exist in order to achieve the design objective. Even under the damage-tolerance method discussed in Section H25.2 of this Appendix, it is the general practice within industry to conduct damage-tolerance tests for design information and guidance purposes. Damage location and growth data must be considered in establishing a recommended inspection program.
Assessing the fatigue characteristics of certain structural elements, such as major fittings, joints, typical skin units, and splices, to ensure that the anticipated service life can reasonably be attained, is required for structure to be evaluated under Section 25.571(c).
(a) Typical loading spectra expected in service. The loading spectrum must be based on measured statistical data of the type derived from government and industry load history studies and, where insufficient data are available, on a conservative estimate of the anticipated use of the airplane. The principal loads to be considered in establishing a loading spectrum are as follows:
(1) Flight loads, gust and maneuver.
(2) Ground loads, taxiing, landing impact, turning, engine runup, braking, and towing.
(3) Pressurization loads.
The development of the loading spectrum requires the definition of the expected flight plan which involves climb, cruise, descent, flight times, operational speeds and altitudes, and the approximate time to be spent in each of the operating regimen. Operations for crew training, and other pertinent factors, such as the dynamic stress characteristics of any flexible structure excited by turbulence, must also be considered. For pressurized cabins, the loading spectrum must, under Section 25.571(b)(5), include the repeated application of the normal operating pressure, and the superimposed effects of flight loads and external aerodynamic pressures.
(b) Components to be evaluated. In assessing the possibility of serious fatigue failures, the design must be examined to determine probable points of failure in service. In this examination, consideration must be given, as necessary, to the results of stress analyses, static tests, fatigue tests, strain gage surveys, tests of similar structural configurations, and service experience. Service experience has shown that special attention must be focused on the design details of important discontinuities, main attach fittings, tension joints and splices, and cutouts such as windows, doors, and other openings. Locations prone to accidental damage (such as that due to impact with ground servicing equipment near airplane doors) or to corrosion must also be considered.
(c) Analyses and tests. Unless it is determined from the foregoing examination that the normal operating stresses at critical regions in the structure are of such a low order that the probability of serious damage growth is extremely low, repeated load analyses or tests must be conducted on structure representative of components or subcomponents of the wing, control surfaces, empennage, fuselage, landing gear, and their related primary attachments. Test specimens must include structure representative of attachment fittings, major joints, changes in section, tensile area cutouts, and discontinuities. Any method used in the analyses must be supported, as necessary, by test or service experience.
H25.2 Damage-tolerance (fail-safe) evaluation. -- (a) General. The damage-tolerance evaluation of structure is intended to ensure that should serious fatigue, corrosion, or accidental damage occur within the operational life of the airplane, the remaining structure can withstand reasonable loads without failure or excessive structural deformation until the damage is detected. Included are the considerations historically associated with fail-safe design. The evaluation encompasses establishing the components which are to be designed as damage-tolerant, defining the loading conditions and extent of damage, conducting structural tests or analyses, or both, to substantiate that the design objective has been achieved, and establishing data for inspection programs to ensure detection of damage. This evaluation applies to either single or multiple load path structure. Design features which must be considered in attaining a damage-tolerant structure include the following:
(1) Multiple load path construction and the use of crack stoppers to limit the growth of cracks.
(2) Duplicate structures designed so that a failure occurring in one-half of the structural member will be confined to the failed half and the remaining structure will still possess adequate load-carrying ability.
(3) Backup structure wherein one member carries all the load, with a second member available and capable of assuming the extra load if the primary member fails.
(4) Materials and stress levels that, after initiation of cracks, provide a controlled slow rate of crack propagation combined with high residual strength.
(5) Arrangement of design details to ensure a sufficiently high probability that a failure in any critical structural element will be detected before the strength has been reduced below the level necessary to withstand the loading conditions specified in Section 25.571(b) so as to permit replacement of repair of the failed elements.
(6) Provisions to limit the probability of multiple damage due to fatigue, including the effects of fretting, particularly under high life conditions arising from the following:
(i) A number of small adjacent initial cracks, each of which being less than the minimum detectable length, developing into a long crack.
(ii) Failures, or partial failures, in adjacent areas following an initial failure due to redistribution of loading.
(iii) Simultaneous failure, or partial failure, or multiple load path discrete elements, working at similar stress levels.
Normally, the damage-tolerance assessment consists of a deterministic evaluation of the above design features and this section provides guidelines for this approach. In certain specific instances, however, damage-tolerant design may be more realistically assessed by a probabilistic evaluation employing methods such as risk analysis. They are routinely employed in fail-safe evaluations of airplane systems and have occasionally been used where structure and systems are interrelated. These methods can be of particular value for structure consisting of discrete isolated elements where damage tolerance depends on the ability of the structure to sustain redistributed loads after failures of discrete elements resulting from fatigue, corrosion, or accidental damage. Where considered more appropriate than deterministic analysis, probabilistic analysis may be used if it can be shown that loss of the airplane is extremely improbable and the statistical data employed in the analysis is based on tests or operational experience, or both, of similar structure.
(b) Identification of principal structural elements. Principal structural elements are those which contribute significantly to carrying flight, ground, and pressurization loads, and whose failure could result in catastrophic failure of the airplane. Typical examples of such elements are as follows:
(i) Wing and empennage. (1) Control surfaces, slats, flaps, and their attachment hinges and fittings.
(ii) Integrally stiffened plates.
(iii) Primary fittings.
(iv) Principal splices.
(v) Skin or reinforcement around cutouts or discontinuities.
(vi) Skin-stringer combinations.
(vii) Spar caps.
(viii) Spar webs.
(2) Fuselage. (i) Circumferential frames and adjacent skin.
(ii) Door frames.
(iii) Pilot window posts.
(iv) Pressure bulkheads.
(v) Skin and any single frame or stiffener element around a cutout.
(vi) Skin or skin splices, or both, under circumferential loads.
(vii) Skin or skin splices, or both, under fore-and-aft loads.
(ix) Skin around a cutout.
(x) Skin or stiffener combinations under fore-and-aft loads.
(xi) Window frames.
(c) Extent of damage. Each particular design must be assessed to establish appropriate damage criteria in relation to inspectability and damage-extension characteristics. In any damage determination, including those involving multiple cracks, it is possible to establish the extent of damage in terms of detectability with the inspection techniques to be used, the associated initially detectable crack size, the residual strength capabilities of the structure, and the likely damage-extension rate considering the expected stress redistribution under the repeated stress experience expected in service and with the recommended inspection frequency. Thus, an obvious partial failure must be considered to be the extent of the damage for residual strength assessment, provided a positive determination is made that the fatigue cracks will be detectable by the available inspection techniques at a sufficiently early stage of the crack development. In a pressurized fuselage, an obvious partial failure might be detectable through the inability of the cabin to maintain operating pressure or controlled decompression after occurrence of the damage. The following are typical examples of partial failures to be considered in the evaluation:
(1) Detectable skin cracks emanating from the edge of structural openings or cutouts.
(2) A detectable circumferential or longitudinal skin crack in the basic fuselage structure.
(3) Complete severance of interior frame elements or stiffeners in addition to a detectable crack in the adjacent skin.
(4) A detectable failure of one element where dual construction is utilized in components such as spar caps, window posts, window or door frames, and skin structure.
(5) The presence of a detectable fatigue failure in at least the tension portion of the spar web or similar element.
(6) The detectable failure of a primary attachment, including a control surface hinge and fitting.
(d) Inaccessible areas. In cases where inaccessible and uninspectable blind areas exist, and suitable damage tolerance cannot practically be provided to allow for extension of damage into detectable areas, the structure must be shown to comply with the fatigue (safe-life) requirements in order to ensure its continued airworthiness. Inspectable structures are generally qualified to the damage-tolerance provisions.
(e) Testing of principal structural elements. The nature and extent of tests on complete structures or on portions of the primary structure will depend upon applicable previous design, construction, tests, and service experience, in connection with similar structures. Simulated cracks must be as representative as possible of actual fatigue damage. Where it is not practical to produce actual fatigue cracks, damage may be simulated by cuts made with a fine saw, sharp blade, guillotine, or other suitable means. In those cases where it is necessary to simulate damage at joints or fittings, bolts may be removed to simulate the failure if this conditions would be representative of an actual failure.
(f) Identification of locations to be evaluated. The locations of damage to structure for damage-tolerance evaluation must be identified as follows:
(1) Determination of general damage locations. The location and modes of damage may be determined by analysis or by fatigue tests on complete structures or subcomponents. However, tests may be necessary when the basis for analytical prediction is not reliable, such as for complex components. If less than the complete structure is tested, care must be taken to ensure that the internal loads and boundary conditions are valid. If a determination is made by analysis, factors such as the following are to be taken into account:
(i) Strain data on undamaged structure to establish points of high stress concentration as well as the magnitude of the concentration.
(ii) Locations where permanent deformation occurred in static tests.
(iii) Locations of potential fatigue damage identified by fatigue analysis.
(iv) Design details which service experience or similarly designed components indicate are prone to fatigue or other damage. In addition, the areas of probable damage from sources such as severe corrosive environment must be determined from a review of the design and past service experience.
(2) Selection of critical damage areas. The process of actually locating where damage must be simulated in principal structural elements identified in paragraph (b) of this section requires taking into account factors such as the following:
(i) Review analysis to locate areas of maximum stress and low margin of safety.
(ii) Selecting locations in an element where the stresses in adjacent elements would be the maximum with the damage present.
(iii) Selecting partial fracture locations in an element where high stress concentrations are present in the residual structure.
(iv) Selecting locations where detection would be difficult.
(g) Damage-tolerance analysis and tests. It is to be determined by analysis, supported by test evidence, that the structure with the extent of damage established for residual strength evaluation can withstand the specified design limit loads (considered as ultimate loads), and that the damage growth rate under the repeated loads expected in service (between the time at which the damage becomes initially detectable and the time at which the extent of damage reaches the value for residual strength evaluation) provides a practical basis for development of the inspection program and procedures described in paragraph (h) of this section. The repeated loads must be as defined in the loading, temperature, and humidity spectra. The loading conditions must take into account the effects of structural flexibility and rate of loading where they are significant,. The damage-tolerance characteristics may be shown analytically by reliable or conservative methods such as the following:
(1) By demonstrating quantitative relationships with structure already verified as damage tolerant.
(2) By demonstrating that the damage would be detected before the damage extent for residual strength evaluation is reached.
(3) By demonstrating that the repeated load stresses do not exceed those of previously verified designs of equivalent materials and inspectability.
The maximum extent of immediately obvious damage from discrete sources must be determined and the remaining structure shown to have static strength for the maximum load (considered as ultimate load) expected during the completion of the flight. Normally this would be an analytical assessment. In the case of uncontained engine failures, the fragments and paths to be considered must be consistent with those used in showing compliance with Section 25.903(d)(1) and with typical damage experienced in service.
(h) Inspection. Detection of damage before it becomes dangerous is the ultimate control in ensuring the damage-tolerance characteristics of the structure. Therefore, the applicant must provide sufficient guidance information to assist operators in establishing the frequency, extent, and methods of inspection of the critical structure. Due to the inherent complex interactions of the many parameters affecting damage tolerance, such as operating practices, environmental effects, load sequence on crack growth, and variations in inspection methods, related operational experience must be taken into account in establishing inspection procedures. Comparative analysis may be used to guide the changes from successful past practice when necessary. Therefore, maintenance and inspection requirements must recognize the dependency on experience and must be specified in a document that provides for revision as a result of operational experience, such as the one containing the manufacturer's recommended structural inspection program.

H25.3 Fatigue (safe-life) evaluation. -- (a) General. The evaluation of structure under the following fatigue (safe-life) strength evaluation methods is intended to ensure that the probability of a catastrophic fatigue failure, as a result of the repeated loads of variable magnitude expected in service, is extremely low throughout its operational life. Under these methods, loading spectra are established, the fatigue life of the structure for the spectra is determined, and a scatter factor is applied to the fatigue life to establish the safe-life for the structure. The evaluation must include consideration of the following:
(1) Estimating or measuring the expected loading spectra for the structure.
(2) Conducting a structural analysis including consideration of the stress concentration effects.
(3) Fatigue testing of structure which cannot be related to a test background to establish response to the typical loading spectrum expected in service.
(4) Determining reliable replacement times by interpreting the loading history, variable load analyses, fatigue test data, service experience, and fatigue analyses.
(5) Providing data for inspection and maintenance instructions and guidance information to the operators to prevent catastrophic failure.
In some instances, it may be necessary to correlate the loadings used in the analysis with flight load and strain surveys.
(b) Safe-life determination: scatter factor. In the interpretation of fatigue analyses and test data, the effect of variability must be accounted for by an appropriate scatter factor. Relating test results to the recommended safe-life is extremely difficult since there are a number of considerations peculiar to each design and test that require evaluation by the applicant. These considerations will depend on the number of representative test specimens, the material, the type of specimen employed, the type of repeated load test, the load levels, and environmental conditions.
(c) Replacement times. Replacement times must be established for parts with established safe-lives and must be included in the information prepared under Section 25.1529. These replacement times may be extended only if additional data indicates an extension is warranted. Important factors to be considered for such extensions include, but are not limited to, the following:
(1) Comparison of original evaluation with service experience.
(2) Recorded load and stress data. Recorded load and stress data entails instrumenting aircraft in service to obtain a representative sampling of actual loads and stresses experienced. The data to be measured includes -- airspeed, altitude, and load factor versus time data; or airspeed, altitude and strain ranges versus time data; or similar data. This data obtained by instrumenting airplanes in service provides a basis for correlating the estimated loading spectrum with the actual service experience.
(3) Additional analyses and tests. If test data and analyses based on repeated load tests of additional specimens are obtained, a reevaluation of the established safe-life may be made.
(4) Tests of parts removed from service. Repeated load tests of replaced parts may be utilized to reevaluate the established safe-life. The tests must closely simulate service loading conditions. Repeated load testing of parts removed from service is especially useful where recorded load data obtained in service are available since the actual loading experience by the part prior to replacement is known.
(5) Repair or rework of the structure. In some cases, repair or rework of the structure may gain further life.
(d) Type design developments and changes. For design developments, or design changes, involving structural configurations similar to those of a design already shown to comply with the applicable provisions of Section 25.571(c), it may be possible to evaluate the variations in critical portions of the structure on a comparative basis. Typical examples would be redesign of the wing structure for increased loads, and the introduction in pressurized cabins of cutouts having different locations or different shapes, or both. This evaluation involves analysis of the predicted stresses of the redesigned primary structure and correlation of the analysis with the analytical and test results used in showing compliance of the original design with Section 25.571(c).

(Secs. 313(a), 601, and 603, Federal Aviation Act of 1958, as amended (49 U.S.C. 1354(a), 1421, and 1423); sec. 6(c), Department of Transportation Act (49 U.S.C. 1655(c); 14 CFR 11.45.)

Note. -- The Federal Aviation Administration has determined that this document does not contain a major proposal requiring preparation of an Economic Impact Statement under Executive Order 11821, as amended, Executive Order 11949, and OMB: Circular A-107.


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Issued in Washington, D.C., on August 9, 1977.
R. P. Skully
Director, Flight Standards Service.
[FR Doc. 77-23420 Filed 8-12-77; 8:45 am]


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Other Notice of Proposed Rulemaking Actions:
Not Applicable.

Final Rule Actions:
Final Rule. Docket No. 16280; Issued on 09/28/78.