3 Identify the structure to a level of detail required to support post-test activities that the applicant will use to evaluate the residual-strength capabilities of the structure.
Structure is free from WFD if the residual strength meets or exceeds that required by § 25.571(b). Therefore, the post-test activities, such as teardown inspections and residualstrength tests, must provide data that support the determination of strength.
● For teardown inspections, specific structural details (e.g., holes, radii, fillets, cutouts) need to be identified.
● For residual-strength testing, the identification at the component or subcomponent level (e.g., longitudinal skin splices) may be sufficient.
(c) Step 3 – Evaluation of WFD-Susceptible Structure. Applicants must evaluate all susceptible structure identified in Step 2. Applicants must demonstrate, by full-scale fatigue-test evidence, that WFD will not occur in the airplane structure prior to the LOV. This demonstration typically entails full-scale fatigue testing, followed by teardown inspections and a quantitative evaluation of any finding or residual-strength testing, or both. Additional guidance about full-scale fatigue-test evidence is included in appendix 2 of this AC.
(d) Step 4 – Finalize LOV. After all susceptible structure has been evaluated, finalize the LOV. The results of the evaluations performed in Step 3 will either demonstrate that the strength at the candidate LOV meets or exceeds the levels required by § 25.571(b), or not. If it is demonstrated that the strength is equal to or greater than that required, the final LOV could be set to the candidate LOV without further showing. If it is demonstrated that the strength is less than the required level, at least two outcomes are possible.
1 Final LOV may equal the candidate LOV. However, this would result in maintenance actions, design changes, or both maintenance actions and design changes to support operation of airplanes up to LOV. For MSD/MED, an applicant may use damagetolerance-based inspections to supplement replacement or modification required to preclude WFD when those inspections have been shown to be practical and reliable.
(aa) Maintenance actions. In some cases, maintenance actions may be necessary for an airplane to reach its LOV. These maintenance actions could include inspections, modifications, replacements, or any combination of these. The applicant must substantiate the maintenance actions per the guidance contained in this AC.
For initial certification, these actions should be both specified as airworthiness-limitation items and incorporated into the ALS of the ICA.
For post-certification airplanes, these actions should be specified as service information by the TC holder and may be mandated by FAA airworthiness directives. Refer to AC 120-104 for addi tional maintenance-action guidance.
(bb) Design changes. The applicant may determine that developing design changes to prevent WFD in future production airplanes is to their advantage.
The applicant must substantiate the design changes per the guidance contained in this AC.
- 2 Final LOV may be less than the candidate LOV. This could preclude having maintenances actions or making design changes.
(aa) In addition to the technical considerations, the LOV may be based on several other factors, including— Maintenance considerations.
Operator input.
Economics.
d. Airworthiness Limitations section (ALS). In accordance with 14 CFR 21.50, the type-certificate holder must provide the ICA, which includes the ALS, with the airplane.
However, the type-certificate holder may or may not have completed the full-scale fatigue-test program at the time of type certification.
(1) Fatigue testing is not completed. Under 14 CFR 25.571, the FAA may issue a type certificate for an airplane model prior to the applicant’s completion of full-scale fatigue testing, provided that the Administrator has approved a plan for completing the required tests.
Until the full-scale fatigue testing is completed and the FAA has approved the LOV, the typecertificate holder must establish a limitation that is equal to one-half the number of cycles accumulated on the test article. (See appendix 2 in this AC for more information about the test article.) At the time of type certification, the applicant should show that at least one calendar year of safe operation has been substantiated using fatigue-test evidence. Under appendix H to part 25, the ALS must contain a defined limitation equal to one-half the number of cycles accumulated on the fatigue-test article approved under § 25.571. This limitation is an airworthiness limitation. No airplane may be operated beyond this limitation until the fatigue testing is completed and an LOV is approved. As additional cycles on the fatigue-test article are accumulated, this limitation may be adjusted accordingly. Upon completion of the full-scale fatigue test, applicants should perform specific inspections and analyses to determine whether WFD has occurred. Additional guidance about post-test WFD evaluations is included in appendix 2 of this AC.
(2) Fatigue testing is completed. After the full-scale fatigue test has been completed, the applicant must include the following in the ALS (see appendix 2 of this AC for additional guidance):
Under appendix H to part 25, the ALS must contain the LOV stated as a number of total accumulated flight cycles or flight hours approved under § 25.571.
Depending on the results of the evaluation under Step 3, the ALS may also include requirements to inspect, modify, or replace structure.
e. Type-design changes (Amendment 25-96 or later). Any person applying for an amended type certificate or a supplemental type certificate to introduce a type design change that adds or affects WFD-susceptible structure must demonstrate that new and affected structure are 14
free from WFD up to the LOV of the airplane. This demonstration may include analysis supported by fatigue-test evidence (see appendix 2 of this AC), analysis that correlates to relevant existing full-scale fatigue-test results, or both. Where analysis supported by test evidence demonstrates freedom from WFD, the applicant may use one of the following approaches to demonstrate freedom from WFD. One approach is based on fatigue-crack initiation and relies on fatigue data. Another approach is based on crack growth and requires the application of fracture-mechanics methods. Inspections or other procedures necessary to prevent catastrophic failure from accidental, environmental, or fatigue damage (including WFD) up to the LOV must be included in the ALS of the ICA, as required by § 25.1529. (See also appendix H to part 25). As an alternative for accidental and environmental damage, the documents containing the ICA may be referenced in the ALS of the ICA. See section 6i of this AC for further guidance.
f. Repairs (Amendment 25-96 or later ). Any person performing a major repair that adds or affects WFD-susceptible structure must demonstrate that any new and affected structure is free from WFD up to the LOV of the airplane. If a repair does not add or affect WFD- susceptible structure, no further WFD evaluation is required. Where analysis supported by test evidence demonstrates freedom from WFD, the applicant may use one of the following approaches to demonstrate freedom from WFD. One approach is based on fatigue-crack initiation and relies on fatigue data. Another approach is based on crack growth and requires the application of fracture-mechanics methods. Refer to appendix 5 of AC 120-93 for the repairapproval process. If WFD is determined likely to occur before the LOV is reached, the applicant must either—
(1) redesign the proposed repair to preclude WFD from occurring before the airplane reaches the LOV, or
(2) develop maintenance actions to preclude WFD from occurring before the airplane reaches the LOV. For repairs, an applicant must identify and include these actions as part of the repair data.
g. Extended LOV. If an applicant proposes to extend an LOV, they must comply with the requirements of 14 CFR 26.23. Refer to AC 120-104 for guidance on extending an LOV.
Typically, the data necessary to extend an LOV includes additional full-scale fatigue-test evidence. The primary source of this test eviden ce should be full-scale fatigue testing. This testing should follow the guidance contained in appendix 2 of this AC.
8. Fatigue Evaluation.
a. General. 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 – will be avoided throughout the structure's operational life. Under these methods, the fatigue life of the structure should be determined. The evaluation should include 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) Performing fatigue testing of structure that cannot be evaluated, based on previous testing, 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) Evaluating the possibility of fatigue initiation from sources, such as stress corrosion, disbonding, environmental (e.g., corrosion) and accidental damage, and manufacturing defects, based on a review of the design, quality control, and past service experience; and
(6) Providing necessary maintenance programs and replacement times to the operators. The maintenance program should be in cluded in the ICA in accordance with 14 CFR
25.1529.
b. Scatter Factor for Safe-life Determination. In the interpretation of fatigue analyses and test data under § 25.571(c), the effect of variability should be accounted for by an appropriate scatter factor. The applicant should ju stify the scatter factor chosen for any safe-life part. The following guidance is provided (See Figure 1):
(1) The base scatter factors applicable to test results are BSF1= 3.0 and BSF2 ≥ 3.0 (see section 8b(5) of this AC). If the a pplicant can meet the criteria of section 8b(3) of this AC, they may use BSF1. As an option, the applicant may elect to use BSF 2. If the applicant cannot meet the criteria of section 8b(3) of this AC, they must use BSF2.
(2) Base scatter factor BSF1 development. The base scatter factor BSF1 is developed using test results of one representative test specimen.
(3) Justification for use of BSF1. BSF 1 may be used only if the following criteria are met:
(a) Understanding of load paths and failure modes. Service and test experience of similar in-service components that were designed using similar design criteria and methods should demonstrate that the load paths and potential failure modes of the components are well understood.
(b) Control of design, material, and manufacturing process quality. The applicant should demonstrate that his quality system (e.g., design, process control, and material standards) ensures that the scatter in fatigue properties is controlled and that the design of the fatigue-critical areas of the part account for the material scatter.
(c) Representativeness of the test specimen.
1 The test article should be full-scale (component or sub-component) and represent that portion of the production aircraft requiring test. All differences between the 16
test article and production article should be accounted for either by analysis supported by test evidence or by testing itself.
2 Construction details, such as bracket attachments and clips, should be accounted for, even though the items themselves may be non-load bearing.
3 Points of load application and reaction should accurately reflect those of the aircraft, ensure correct behavior of the test article, and guard against uncharacteristic failures.
4 Systems used to protect the structure against environmental degradation can have a negative effect on fatigue life and, therefore, should be included as part of the test article.
(4) Adjustments to base scatter factor BSF1. If the criteria in section 8b(3) justifying the use of BSF1 have been met, the base value of 3.0 should be adjusted to account for the following factors, where these factors are not wholly taken into account by design analysis.
As a result of the adjustments, the final scatter factor may be less than, equal to, or greater than
3.0.
(a) Material fatigue scatter. Material properties should be investigated up to a 99% probability of survival and a 95% level of confidence.
(b) Spectrum severity. The test-loads spectrum should be derived, based on a spectrum-sensitive analysis accounting for variations in utilization (e.g., aircraft weight and center of gravity), as well as for the occurrence and size of loads. The test-loads spectrum applied to the structure should be demonstrated to be conservative when compared to the usage expected in service.
(c) Number of representative test specimens. Well established statistical methods should be used that associate the number of items tested with the distribution chosen to obtain an adjustment to the base scatter factor.
(5) Adjustments to base scatter factor BSF2. If section 8b(3) of this AC cannot be satisfied in its entirety, the applicant should use BSF2.
(a) The applicant should propose scatter factor BSF2 based on careful consideration of the following issues:
● required level of safety, ● number of representative test specimens, ● how representative the test is, ● expected fatigue scatter, 17
HAVE THE CRITERIA OF§ 8b(3) BEEN MET:
-- Service and Text Experience of Similar Components § 8b(3)(a) -- QA System Ensuring Fatigure Scatter Lies within Certain Limits § 8b(3)(b) --Representativeness of Test Specimen§ 8b(3)(c)
ALL CRITERIA MET USE BSF1 = 3.0 § 8b(3) HAVE THE ELEMENTS OF§ 8b(4) BEEN ACCOUNTED FOR IN DESIGN:
-- Fatigue Scatter to Account for P=99% and C=95% § 8b(4)(a) -- Spectrum Severity § 8b(4)(b)
ALL ELEMENTS MET ADJUST BSF1 for:
-- Number of Specimen Tested § 8b(4)(c) -- Resolution of Test Loads to Actual Loads § 8b(6)
SOME ELEMENTS MISSED ADJUST BSF1 for:
-- Fatigue Scatter§ 8b(4)(a) -- Spectrum Severity§ 8b(4)(b) -- Number of Specimen Tested § 8b(4)(c) -- Resolution of Test Loads to Actual Loads § 8b(6) SAFE LIFE= TEST CYCLES/SCATTER FACTOR*
SOME CRITERIA MISSED USE BSF2 ~ 3.0 BSF2 DETERMINED FROM ANALYSIS AND TEST § 8b(5)(a) -- Required Level of Safety -- Number of Specimens Tested -- Representativeness of Test -- Fatigue Scatter to Account for P-99% and C=95% -- Type of Repeated Loads Test -- Accuracy of the Test Load Spectrum -- Spectrum Severity -- Service Environmental Conditions Minimum Value~ 3.0 § Bb(5)(b) Adjust BSF2 for Resolution of Test Loads to Actual Loads § 8b(6) SAFE LIFE= TEST CYCLES/SCATTER FACTOR* *Scatter Factor= BSF,. X Adjustment 1/13/11 AC 25.571-1D Figure 1 – Safe-Life Determination 18
● type of repeated load test, ● accuracy of the test-loads spectrum, ● spectrum severity, and ● expected service environmental conditions.
(b) In no case should the value of BSF2 be less than 3.0.
(6) Resolution of test loadings to actual loadings. The applicant may use a number of different approaches to reduce both the number of load cycles and number of test setups required. These include the following:
● spectrum blocking (i.e., a change in the spectrum load sequence to reduce the total number of test setups);
● high-load clipping (i.e., reduction of the highest spectrum loads to a level at which the beneficial effects of compression yield are reduced or eliminated); and ● low-load truncation (i.e., the removal of non-damaging load cycles to simplify the spectrum).
(7) Due to the modifications to the flight-by-flight loading sequence caused by these changes, the applicant should propose either analytical or empirical approaches to quantify an adjustment to the number of test cycles which represents the difference between the test spectrum and the assumed flight-by-flight spectrum. In addition, an adjustment to the number of test cycles may be justified by raising or lowering the test-load levels, as long as data support such adjustment. Other effects to consider are different failure locations, different response to fretting conditions, and temperature effects. The analytical approach should either use wellestablished methods or be supported by test evidence.
c. Replacement times. Under § 25.571(a)(3), replacement times should be established for parts with established safe-lives and should be included in the ALS of the ICA. These replacement times can be extended if additional data indicate an extension is warranted. Factors that should be considered for such extensions include the following:
(1) Comparison of original evaluation with service experience.
(2) Recorded load and stress data. Recorded load and stress data entails the installation of instrumentation on airplanes in service to obtain a representative sampling of actual loads and stresses experienced. The data to be measured include airspeed, altitude, and load-factor-versus-time data; or airspeed, altitude, and strain-ranges-versus-time data; or similar data. The data obtained from airplanes in serv ice provide a basis for correlating the estimated loading spectrum with the actual service experience.
(3) Additional analyses and tests. If te st data and analyses are obtained based on repeated load tests of additional specimens, a re-evaluation of the established safe-life can be made.
(4) Tests of parts removed from service. Repeated load tests of replaced parts can be used to re-evaluate the established safe-life. The tests should closely simulate service loading conditions. Repeated load testing of pa rts removed from service is especially useful where recorded load data obtained in service are available, since the actual loading experienced by the part prior to replacement is known.
(5) Repair or rework of the structure. In some cases, repair or rework of the structure can 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 § 25.571(c), it may be possible to evaluate the variations in critical portions of the structure on a comparative basis. Examples would be (1) redesign of the wing structure for increased loads or (2) the introduction, in pressurized cabins, of cutouts having different locations or different shapes, or both. This evaluation should involve 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 with § 25.571(c) of the original design.
e. Environmental effects such as temperature and humidity should be considered in the damage-tolerance and fatigue analysis, and should be demonstrated through suitable testing.
9. Discrete Source Damage.
a. General. The purpose of this section is to establish FAA guidelines for consistent selection of load conditions for residual-strength substantiation in showing compliance with § 25.571(e), Damage-tolerance (discrete source) evaluation. The intent of these guidelines is to define, with a satisfactory level of confidence, load conditions that will not be exceeded 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 pilot knowledge that the airplane may be in a damaged state. With these considerations in mind, use the following ultimate loading conditions to establish residual strength of the damaged structure.
b. 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:
1/1113/11 AC 25.S71HDD
(a) The maximum, normall, operating diflerenttahl pressune, multiplodd by a
1.1 factor, plus the expected externall aerodynamicc pressuress during; 1g level flight combinedd with Ig flight loads.
(b) The airplane, assumed to be in 1g level flight, should be showm to be able to survive any maneuverr or any other flight-path deviation caused by the specified incidentt of § 25.571(e), taking into accountt any likely damage to the flight controlls and pilot nonnall comeccitive action.
(2) Following: the incidentt:
(a) Seventy percent (70%) limit flightmanouveer loads and, separatedly, 40% of the limitt gust velocity (verticall and lateral) at the specified speeds, each combinedd with the maximum appropriate cabim differenttall pressune (includingg the expectad exterall aerodynamicc pressure)).
(b) The airplane must be showm by anallysis to be free from flutter up to Vo/Mo with any chamge in structunall stiffiness resulting from the incidentt.
Jiffing E. Dum.
Jeffrey E. Duvem Acting Mamager;, Tramsportt Airplame Directoratte Aircraft Certification Service 21
Appendix 1 Appendix 1 References and Definitions
1. References.
a. AC 20-107B, “Composite Aircraft Structure”
b. AC 25.1529-1A, “Instructions for Continued Airworthiness of Structural Repairs on Transport Airplanes”
c. AC 91-56B, “Continuing Structural Integrity Program for Airplanes”
d. AC 120-93, “Damage Tolerance Inspections for Repairs and Alterations”
e. AC 120-104, “Widespread Fatigue Damage and Establishing LOV”
2. Definitions of Terms Used in this AC.
a. Airworthiness Limitation item (ALI) — A mandatory-maintenance action identified in the Airworthiness Limitations section of a design-approval holder’s Instructions for Continued Airworthiness. These items may co ntain mandatory modification or replacement times, mandatory inspection thresholds, intervals, and inspection procedures.
b. Airworthiness Limitations section (ALS) — Relative to the Instructions for Continued Airworthiness, the ALS is a collection of mandatory-maintenance actions required for airplane structure and fuel-tank systems. For structural-maintenance actions, the ALS includes structural-modification times, structural-replacement times, structural-inspection thresholds and intervals, and structural-inspection procedures.
c. Alteration or modification — A design change that is made to an airplane. Within the context of this AC, the two terms are synonymous.
d. Damage tolerance — The attribute of the structure that permits it to retain its required residual strength for a period of use after the structure has sustained a given level of fatigue, corrosion, or accidental or discrete source damage.
e. Design service goal — The period of time (in flight cycles or flight hours, or both) established at design and/or certification during which the airplane structure is reasonably free from significant cracking.
f. Fail-safe — The attribute of the structure that permits it to retain its required residual strength for a period of unrepaired use after the failure or partial failure of a principal structural element.
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Appendix 1
g. Instructions for Continued Airworthiness (ICA) — Documentation that sets forth instructions and requirements for the maintenance that is essential to the continued airworthiness of an aircraft, engine, or propeller.
h. Multiple load path — Applies to structure, the applied loads of which are distributed through redundant structural members, so that the failure of a single structural member does not result in the loss of structural capability to carry the applied loads.
i. Limit of validity (of the engineering data that supports the structural maintenance program) — The period of time (in flight cycles, flight hours, or both), up to which it has been demonstrated by test evidence, analysis and, if available, service experience and teardown inspection results of high-time airplanes, that widespread fatigue damage will not occur in the airplane structure.
j. Principal structural element (PSE) — An element that contributes significantly to the carrying of flight, ground, or pressurization loads, and whose integrity is essential in maintaining the overall structural integrity of the airplane. Principal structural elements include all structure susceptible to fatigue cracking, which could contribute to a catastrophic failure.
Refer to appendix 5 of this AC for clarification on how this term relates to the terms, fatigue critical structure (FCS) and WFD – susceptible structure.
k. Safe-life — The number of events, such as flight cycles, landings, or flight hours, within which the structure strength has a low probability of degrading below its design ultimate value due to fatigue cracking.
l. Scatter factor — A life-reduction factor used in interpreting fatigue analysis and test results.
m. Single load path — Describes structure, the applied loads of which are eventually distributed through a single structural member, the failure of which would result in the loss of the structural capability to carry the applied loads.
n. Teardown inspection — The term used for the process of disassembling structure and using destructive inspection techniques or visual (e.g., magnifying glass and dye penetrant) or other non-destructive (e.g., eddy current, ultrasound) inspection techniques to identify the extent of damage, within a structure, caused by fatigue, corrosion, and accidental damage.
o. Widespread fatigue damage (WFD) — The simultaneous presence of cracks at multiple structural locations that are of sufficient size and density such that the structure will no longer meet the residual-strength requirements of § 25.571(b).
(1) Multiple site damage (MSD) — A source of widespread fatigue damage characterized by the simultaneous presence of fatigue cracks in the same structural element.
(2) Multiple element damage (MED) — A source of widespread fatigue damage characterized by the simultaneous presence of fatigue cracks in adjacent structural elements.
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Appendix 1
(3) Structural modification point (SMP) — The point in time when a structural area must be modified to preclude WFD.
(4) Inspection start point — The point in time when special inspections of the fleet are initiated due to a specific probability of having a MSD/MED condition.
p. WFD (average behavior) — The point in time, without intervention, when 50% of the fleet is expected to develop WFD for a particular structure.
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