# AC 25.571-1A: Damage-Tolerance and Fatigue Evaluation of Structure

- Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it)
- Status: Historical
- Version in force: [AC 25.571-1D](https://openairdocs.org/ac/25.571-1), approved January 13, 2011. This is a historical version
- Approved: March 5, 1986
- Last reviewed: September 8, 2014
- Family: AC 25.571-1
- Cancels: [AC 25.571-1](https://openairdocs.org/ac/25.571-1)
- Cancelled by: [AC 25.571-1B](https://openairdocs.org/ac/25.571-1B)
- 14 CFR parts: 25, 121
- Pages: 12
- PDF: https://openairdocs.org/files/sha256/e7/13/e713dafa7d317a26acf7e63472557150efb702aa3b04c03abd3e251f1864bad8
- FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/AC83B2D952537F6986257D4D0065ED5A.0001
- DRS record last changed: November 21, 2025
- Page: https://openairdocs.org/ac/25.571-1A

## Versions

- [AC 25.571-1D](https://openairdocs.org/ac/25.571-1) (Current), approved Jan 13, 2011, 41 pages
- [AC 25.571-1C](https://openairdocs.org/ac/25.571-1C) (Historical), approved Apr 29, 1998, 20 pages
- [AC 25.571-1B](https://openairdocs.org/ac/25.571-1B) (Historical), approved Feb 18, 1997, 17 pages
- [AC 25.571-1A](https://openairdocs.org/ac/25.571-1A) (Historical), approved Mar 5, 1986, 12 pages
- [AC 25.571-1](https://openairdocs.org/ac/25.571-1?v=FAC94DCE61E6AFE286257D4D0065B95E) (Historical), approved Sep 28, 1978, 12 pages

## ADs that cite it (4)

- [AD 2019-23-01](https://openairdocs.org/ad/2019-23-01) (historical): Time Limits/ Maintenance Checks · Airbus A318-111, A318-112, A318-121 and 37 more. Effective Jan 9, 2020
- [AD 2019-23-02](https://openairdocs.org/ad/2019-23-02) (historical): Time Limits/ Maintenance Checks · Airbus A330-201, A330-202, A330-203 and 13 more. Effective Dec 30, 2019
- [AD 2004-07-22](https://openairdocs.org/ad/2004-07-22) (historical): Structural elements · Boeing 747 Series (all). Effective May 12, 2004
- [AD 2002-19-14](https://openairdocs.org/ad/2002-19-14): Fastener holes in the left and right side lower skin panels · Hawkins & Powers C-130A, C-130A, C-130A and 4 more. Effective Sep 26, 2002

## Text

Each page's text from the PDF; open the PDF at a page with #page=N. Some pages are OCR from the scanned FAA advisory circular. OCR can misread letters and numbers. Those pages are marked (OCR).

### Page 1 (OCR)

Advisory U.S. Department of Transportation Circular Federal Aviation Administration Subject: DAMAGE-TOLERANCE AND FATIGUE Date: 3/5/86 AC No: 25.571-1A EVALUATION OF STRUCTURE Initiated by: ANM-110 Change:

PURPOSE. This advisory circular (AC) sets forth an acceptable means of compliance with the provisions of Part 25 of the Federal Aviation Regulations (FAR) dealing with the damage-tolerance and fatigue evaluation requirements of aircraft structure.

2. CANCELLATION. Advisory Circular 25.571-1, dated 9/28/78, is cancelled.

**DEFINITIONS OF TERMS USED IN THIS AC.**

a . Damage tolerance means that the structure has been evaluated 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.

b. Fail-safe means that the structure has been evaluated to assure that catastrophic failure is not probable after fatigue failure or obvious partial failure of a single, principal structural element.

Safe-life means that the structure has been evaluated to be able to withstand the repeated loads of variable magnitude expected during its service life without detectable cracks.

d. Principal structural elements are those which contribute significantly see it it cantly to carrying flight, ground, and pressurization loads, and whose failure could result in catastrophic failure of the airplane.

Critical structural elements are those elements whose failure would result in catastrophic failure of the airplane.

f. Primary structure is that structure which carries flight, ground, or pressure loads.

g. Secondary structure is that structure which carries only air or inertial loads generated on or within the secondary structure.

'h. Single load path is where the applied loads are eventually distributed through a single member within an assembly, the failure of which would result in the loss of the structural integrity of the component involved.

### Page 2 (OCR)

AC 25.571-1A 3/5/86

i. Multiple load path is identified with redundant structures in which, (with the failure of individual elements) the applied loads would be safely distributed to other load carrying members.

history has demonstrated to be reliable. j. Reliability refers to detail designs or methodologies which service

k. Probability refers to a probability of occurrence of an event consistent with past successful experience.

**4. BACKGROUND.**

a. 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 the existing fatigue requirements in §$ 25.571 and 25.573 of Part 25 of the FAR, 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 (41 FR 50956). The proposals and related discussions formed the basis for the revision of the structural fatigue evaluation standards of §$ 25.571 and 25.573 § 25.573 was deleted (the scope of § 25.571 was expanded to cover the substance and the development of guidance material. To that end, $ 25.571 was revised, of the deleted section), and guidance material (AC 25.571-1) containing compliance provisions related to the proposed change was provided.

including discrete source damage, has been developed and is incorporated in this revision. Since issuance of AC 25.571-1 on 9/28/78, additional guidance material,

**5. INTRODUCTION.**

The contents of this advisory circular are considered by the FAA in determining compliance with the damage-tolerance and fatigue requirements of $ 25.571.

(1) Although a uniform approach to the evaluation required by § 25.571 is desirable, it is recognized that in such a complex field new design features and methods of fabrication, new approaches to the evaluation, and new configurations could necessitate variations and deviations from the procedures described in this advisory circular. Close adherence to the procedures contained in this advisory circular should be encouraged.

(2) Damage-tolerance design is required, unless it entails such complications that an effective damage-tolerant structure cannot be achieved within the limitations of geometry, inspectability, or good design practice.

Under these circumstances, a design that complies with the fatigue evaluation (safe-life) requirements is used. A typical example of structure that might not be conducive to damage-tolerance design is the landing gear and its attachments.

(3) Experience with the application of methods of fatigue evaluation indicates that a test background should exist in order to achieve the desig Page 2 Par 3

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objective. Even under the damage-tolerance method discussed in paragraph 6 of this AC, it is the general practice within industry to conduct damage-tolerance tests for design information and guidance purposes. Damage location, growth, and detection data should also be considered in establishing a recommended Typical loading spectrum expected in service. The loading spectrum should 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 that should be considered in establishing a loading spectrum are flight loads (gust and maneuver), ground loads (taxiing, landing impact, turning, engine runup, braking, thrust reversing, and tovirun includes the ressurization loads. The development of the loading spectrum 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 regimes. Operations for crew training and other pertinent factors, such as the dynamic stress characteristics of any flexible structure excited by turbulence, should also be considered. For pressurized cabins, the loading spectrum should include the repeated application of the normal operating differential pressure, and the superimposed effects of flight loads and external aerodynamic pressures.

Components to be evaluated. In assessing the possibility of serious fatigue failures, the design should be examined to determine probable points of failure in service. In this examination, consideration should 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 should be focused on the design details of important discontinuities, main attach Open servicing anuipment hear streline dong but coring dus no impact with openings. Locations Fittings, tension joints, splices, considered.

Analyses and tests. Unless it is determined from the foregoing examination that the normal operating stresses in specific regions of the structure are of such a low order that serious damage growth is extremely improbable, repeated load analyses or tests should be conducted on structures represea fuself e, panding ora, comptits ef ted primary attachments. fest specimens should include structure representative of attachment fittings, major joints, changes in section, cutouts, and discontinuities. Any method used in the analyses should be supported, as necessary, by test or service experience.

Typical (average) values of material properties and other parameters may be used in residual strength, crack growth, and damage detection analyses for damage tolerance evaluations per paragraph 6 and discrete source damage per paragraph 8.

**6. DAMAGE-TOLERANCE EVALUATION.**

a . General. The damage-tolerance evaluation of structure is intended to ensure that should serious fatigue, corrosion, or accidental damage occur within Par 5 Page 3

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AC 25.571-1A 3/5/86 damage is detected. with fail-safe design. The evaluation should encompass establishing the the operational life of the airplane, the remaining structure can withstand reasonable loads without failure or excessive structural deformation until the Included are the considerations historically associated components which are to be designed as damage-tolerant, defining the loading both, to substantiate that the design objective has been achieved, and conditions and extent of damage, conducting structural tests or analyses, or establishing data for inspection programs to ensure detection of damage.

Although this evaluation applies to either single or multiple load path in attaining a damage-tolerant structure include the following: in achieving damage tolerant design. structure, the use of multiple load path structure should be given high priority Design features which should be considered

(1) Multiple load path construction and the use of crack stoppers to control the rate of crack growth, and to provide adequate residual static strength;

provide a controlled slow rate of crack propagation combined with high residual strength; (2) Materials and stress levels that, after initiation of cracks,

(3) 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 § 25.571(b), so as to allow replacement or repair of the failed elements; and

(4) Provisions to limit the probability of concurrent multiple damage, particularly after long service, which could conceivably contribute to a commor fracture path. Examples of such multiple damage are:

single long crack; (i) A number of small cracks which might coalesce to form a redistribution of loading following a failure of a single element; and (ii) Failures, or partial failures, in adjacent areas due to the path discrete elements, working at similar stress levels. (iii) Simultaneous failure, or partial failure, of multiple load design might be more realistically assessed by a probabilistic evaluation for this approach. evaluation of the above design features. b. Normally, the damage tolerance assessment consists of a deterministic In certain specific instances, however, damage-tolerant This paragraph provides guidelines 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 ioads after failures of discrete elements resulting from fatigue, corrosion, or accidental damage. Where considered appropriate on multiple load path structure, probabilistic analysis may be used if it can be show that loss of Page 4 Par 6

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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.

c. Identification of principal structural elements. Principal structural elements are those which contribute sigificantly 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:

(1) Wing and empennage.

(i) Control surfaces, slats, flaps, and their mechanical systems and attachments (hinges, tracks, 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; and

(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;

(viii) Skin around a cutout;

(ix) Skin and stiffener combinations under fore and aft loads;

(x) Door skins, frames, and latches; and

(xi) Window frames.

Par 6 Page 5

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AC 25.571-1A 3/5/86

(3) Landing gear and their attachments.

(4) Engine mounts.

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 initially detectable crack size, the residual strength capabilities of the terms of detectability with the inspection techniques to be used, the associated structure, and the likely damage-extension rate, considering the expected stress redistribution under the repeated loads expected in service and with the expected inspection frequency.

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 operatino pressure or controlled decompression after occurrence of the damage. The the evaluation: following are typical examples of partial failures which should be considered in openings or cutouts; (1) Detectable skin cracks emanating from the edge of structural basic fuselage structure; (2) A detectable circumferential or longitudinal skin crack in the 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 and skin structure; utilized in components such as spar caps, window posts, window or door frames, tension portion of the spar web or similar element; and (5) The presence of a detectable fatigue failure in at least the surface hinge and fitting. (6) The detectable failure of a primary attachment, including a control inspectability of all structural parts, and to qualify them under the damage tolerance provisions (reference Inaccessible areas. Every reasonable effort should be made to ensure $ 25.611).

f. Testing of principal structural elements. The nature and extent of residual strength 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 should be as representative as possible of actual fatigue damage. Where it is not practical to produce actual fatigue cracks, damage can be simulated by cuts Page 6 Par 6

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AC 25.571-1A 3/5/86 nade with a tine saw, sharp blade, guillotine, or other suitable means. If sawcuts in primary structure are used to simulate sharp fatique cracks, sufficient evidence should be available from element tests to indicate equivalent residual strength.

equivalent, is to be simulated as part of a possible damage configuration in joints or fittings, bolts can be removed to provide that part of the simulation.

g. Identification of locations to be evaluated. The locations of damage to structure for damage-tolerance evaluation should be identified as follows:

(1) Determination of general damage locations. The location and modes of damage can be determined by analysis or by fatigue tests on complete structures or subcomponents. However, tests might 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 should be taken to ensure that the internal loads and boundary conditions are valid.

(i) If a determination is made by analysis, factors such as the following should be taken into account:

(A) Strain data on undamaged structure to establish points of high stress concentration, as well as the magnitude of the concentration;

(B) Locations where permanent deformation occurred in static tests;

(C) Locations of potential fatigue damage identified by fatigue analysis; and (D) Design details which service experience of similarly designed components indicates are prone to fatigue or other damage.

in addition, the areas of probable damage from sources such as severe corrosive environment should 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 should be simulated in principal structural elements identified in paragraph 6c of this AC should take into account factors such as the following:

(1) Review analysis to locate areas of maximum stress and low margin of safety;

(fi) Select locations in an element where the stresses in adjacent elements would be the maximum with the damage present;

(iii) Select partial fracture locations in an element where high stress concentrations are present in the residual structure; and

(iv) Select locations where detection would be difficult.

Page 7 Par 6

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AC 25.571-1A 3/5/86

h. Damage-tolerance analysis and tests.

that: (1) It should be determined by analysis, supported by test evidence, residual strength evaluation, can withstand the specified design limit loads (considered as ultimate loads), and (i) The structure, with the extent of damage established for service (between the time the damage becomes initially detectable and the time (ii) The damage growth rate under the repeated loads expected in provides a practical basis for development of the inspection program and procedures described in paragraph 6i of this AC.

(2) The repeated loads should be as defined in the loading, temperature, and humidity spectra. The loading conditions should take into account the effects of structural flexibility and rate of loading where they are significant.

(3) The damage tolerance characteristics can be shown analytically by reliable or conservative methods such as the following:

By demonstrating quantitative relationships with structure already verified as damage tolerant;

reaches the value for residual strength evaluation; or By demonstrating that the damage would be detected before it

(iii) By demonstrating that the repeated loads and limit load stresses do not exceed those of previously verified designs of similar configuration, materials, and inspectability.

(4) The maximum extent of immediately obvious damage from discrete sources should be determined and the remaining structure show to have static completion of the flight. the case of uncontained engine failures, the fragments and paths to be considered should be consistent with those used in showing compliance with strength for the maximum load (considered as ultimate load) expected during the Normally, this would be an analytical assessment. In § 25.903(d) (1) of the FAR, and with typical damage experienced in service.

1. Inspection.

(1) Detection of damage before it becomes critical is the ultimate control in ensuring the damage tolerance characteristics of the structure.

Therefore, the applicant should provide sufficient guidance information to assist operators in establishing the frequency, extent, and methods of inspection of the critical structure. This kind of information must, under § 25.571(a)(3) of the FAR, be included in the instructions for continued Page 8 Par 6

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(2) 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 should be taken into account in establishing inspection procedures.

(3) A comparative analysis can be used to guide the changes from successful past practice when necessary. Therefore, maintenance and inspection requirements should recognize the dependence on experience and should be specified in a document that provides for revision as a result of operational experience, such as the one containing the operator's FAA-approved structural inspection program developed through the Maintenance Review Board (MRB) procedures for FAR Part 121 operators.

**7. FATIGUE EVALUATION.**

The evaluation of structure under the following fatigue (safe-life) strength evaluation methods is intended to ensure that catastrophic fatigue failure, as a result of the repeated loads of variable magnitude expected in service, is extremely improbable throughout the structure's ›perational life. Under these methods, loading spectra should be established, the fatigue life of the structure for the spectra should be determined, and a scatter factor should be applied to the fatigue life to establish the safe-life for the structure. The evaluation should include the following; however, in some instances it might be necessary to correlate the loadings used in the

(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; and

(5) Providing data for inspection and maintenance instructions and guidance information to the operators.

b. Scatter Factor for Safe-life Determination. In the interpretation of fatigue analyses and test data, the effect of variability should, under § 25.571(c), 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 necessitate 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.

Par 6 Page 9

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AC 25.571-1A 3/5/86 following: with established safe-lives and should, under § 25.571(a)(3), be included in the information prepared under § 25.1529. should be considered for such extensions include, but are not limited to, the if additional data indicates an extension is warranted. Important factors which Replacement times. Replacement times should be established for parts These replacement times can be extended

(1) Comparison of original evaluation with service experience.

estimated loading spectrum with the actual service experience. airspeed, altitude, and load factor versus time data; or airspeed, altitude, and instrumenting airplanes in service, provides a basis for correlating the strain ranges versus time data; or similar data. The data, obtained by entails instrumenting airplanes in service to obtain a representative sampling of actual loads and stresses experienced. The data to be measured includes (2) Recorded load and stress data. Recorded load and stress data established safe-life can be made. repeated load tests of additional specimens are obtained, a reevaluation of the (3) Additional analyses and tests. If test data and analyses based on tests should closely simulate service loading conditions. Repeated load testing of parts removed from service is especially useful where recorded load data replaced parts can be utilized to reevaluate the established safe-life. The (4) Tests of parts removed from service. Repeated load tests of obtained in service are available since the actual loading experienced by the part prior to replacement is known.

of the structure can gain further life. (5) Repair or rework of the structure. In some cases, repair or rework changes involving structural confiqurations similar to those of a design already shown to comply with the applicable provisions of § 25.571(c), it might 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 involve analysis of the predicted stresses of the redesigned primary structure and correlation of the analysis with the analytical and test results used in different locations or different shapes, or both. This evaluation should showing compliance of the original design with § 25.571(c).

demonstrated through suitable testing. considered in the damage tolerance and fatigue analysis and should be Environmental effects such as temperature and humidity should be

**8. DISCRETE SOURCE DAMAGE.**

the consistent selection of load conditions for residual strength substantiation General. The purpose of this section is to establish FAA guidelines for Page 10 Par 7

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..• in showing compliance with § 25.5/1(e), Damage-tolerance (discrete source) evaluation. The intent of these quidelines is to define load conditions that will not be exceeded with a satisfactory level of confidence on the flight during which the specified incident of $ 25.571 (e) occurs. In defining these load conditions, consideration has been given to the expected damage to the airplane, the anticipated response of the pilot at the time of the incident, and the actions of the pilot to avoid severe load environments for the remainder of the flight consistent with his knowledge that the airplane may be in a damaged With these considerations in mind, the following ultimate loading conditions should be used to establish residual strength of the damaged structure.

The maximum extent of immediately obvious damage from discrete sources ($ 25.571(e)) should be determined and the remaining structure shown, with an acceptable level of confidence, to have static strength for the maximum load (considered as ultimate load) expected during completion of the flight.

c. The ultimate loading conditions should not be less than those developed from the following conditions:

(1) At the time of the incident:

The maximum normal operating differential pressure, multiplied by a 1.1 factor, plus the expected external aerodynamic pressures during 1g level flight, combined with 1 g flight loads.

The airplane, assumed to be in 1 g level flight, should be shown to be able to survive any maneuver or any other flight path deviation caused by the specified incident of $ 25.571 (e), taking into account any likely damage to the flight controls and pilot normal corrective action.

(2) Following the incident:

Seventy percent (70%) limit flight maneuver loads and, separately, 40 percent of the limit gust velocity (vertical or lateral) at the specified speeds, each combined with the maximum appropriate cabin differential pressure (including the expected external aerodynamic pressure).

(11) The airplane must be shown by analysis to be free from flutter up to Vp/Mp with any change in structural stiffness resulting from the incident.

LEROY A. KEITH Manager, Aircraft Certification Division, ANM-100 2 U.S. Governmen 491-810/20706 Par 8 Page 11

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