Annex 1 to Appendix 2 to AMC 20-20A Full-scale fatigue test evidence

• ED Decision 2021/007/R found in: AMC-20 Amdt 23 - Airworthiness of Products, Parts and Appliances (Jan 2022)
• ED Decision 2020/023/R found in: AMC-20 Amdt 21 - Airworthiness of Products, Parts and Appliances (Apr 2021)

Annex 1 to Appendix 2 to AMC 20-20B Full-scale fatigue test evidence ED Decision 2021/007/R Contents (a)       Overview (b)       Full-scale fatigue test evidence (c)       Key elements of a full-scale fatigue test programme (1)       Article (2)       Test set-up and loading (i)        Test set-up (ii)       Test loading (3)       Test duration (i)        New type certificates and derivatives (ii)       Repairs and type design changes (4)       Post-test evaluation (i)        Residual strength tests (ii)       Teardown inspections (d)       Scope of full-scale fatigue test article (1)       New type designs (2)       Derivative models (3)       Type design changes — SBs (4)       Type design changes — STCs (5)       Major repairs (e)       Use of existing full-scale fatigue test data (f)        Use of in-service data (a)       Overview CS 25.571(b) Amendment 19 specifies that special consideration for WFD must be included in the fatigue and damage tolerance evaluation where the design is such that this type of damage could occur. CS 25.571(b) Amendment 19 specifies the effectiveness of the provisions to preclude the possibility of WFD occurring within the limits of validity of the maintenance programme to be demonstrated with sufficient full-scale fatigue test evidence. The determination of what constitutes ‘sufficient full-scale test evidence’ requires a considerable amount of engineering judgement and is a matter that should be discussed and agreed to between an applicant and EASA early in the planning stage of a certification project. Sufficient test evidence is also necessary to support compliance with CS 26.303 and the most straightforward means of compliance is to utilise existing full-scale test evidence. (b)       Full-scale fatigue test evidence In general, sufficient full-scale fatigue test evidence consists of full-scale fatigue testing to at least twice the LOV, followed by specific inspections and analyses to determine that WFD has not occurred. The following factors should be considered in determining the sufficiency of the evidence: Factor 1: The comparability of the load spectrum between the test and the projected usage of the aeroplane. Factor 2: The comparability of the airframe materials, design and build standards between the test article and the certified aeroplane. Factor 3: The extent of post-test teardown inspection, residual strength testing and analysis for determining whether widespread fatigue cracking has occurred. Factor 4: The duration of the fatigue testing. Factor 5: The size and complexity of a design or build standard change. This factor applies to design changes made to a model that has already been certified and for which full-scale fatigue test evidence for the original structure should have already been determined to be sufficient. Small, simple design changes, comparable to the original structure, could be analytically determined to be equivalent to the original structure in their propensity for WFD. In such cases, additional full-scale fatigue test evidence should not be necessary. Factor 6: In the case of major changes and STCs, the age of an aeroplane being modified. This factor applies to aeroplanes that have already accumulated a portion of their LOV prior to being modified. An applicant should demonstrate freedom from WFD up to the LOV in place for the original aeroplane and may take into account the age of the aeroplane being modified. (c)       Key elements of a full-scale fatigue test programme The following guidance addresses key elements of a test programme that is intended to generate the data necessary to support compliance, and it can also be used to evaluate and interpret existing full-scale test data for the purposes of supporting compliance with point 26.303 of Part-26. (1)       Article. The test article should be representative of the structure of the aeroplane to be evaluated (i.e. ideally a production-standard article). The attributes of the type design that could affect MSD/MED initiation, growth and subsequent residual strength capability should be replicated as closely as possible on the test article. Critical attributes include, but are not limited to, the following: —        material types and forms; —        dimensions; —        joining methods and details; —        coating and plating; —        the use of faying surface sealant; —        assembly processes and sequences; and —        the influence of secondary structure (e.g. loads induced due to proximity to the structure under evaluation). (2)       Test set-up and loading. The test set-up and loading should result in a realistic simulation of the expected operational loads. (i)        Test set-up. The test set-up dictates how loads are introduced into the structure and reacted. Every effort should be made to introduce and react loads as realistically as possible. When a compromise is made (e.g. wing air loading), the resulting internal loads should be evaluated (e.g. using finite element methods) to ensure that the structure is not being unrealistically underloaded or overloaded, locally or globally. (ii)       Loading spectrum. The test loading spectrum should include loads from all damaging sources (e.g. cabin pressurisation, manoeuvres, gusts, engine thrust, control surface deflections, and landing impacts) that are significant for the structure being evaluated. Consideration should also be given to temperature and other environmental effects that may affect internal loads. A supporting rationale should be provided when a load source is not represented in a sequence. Additionally, differences between the test sequence and the expected operational sequence should be justified. For example, it is standard practice to eliminate low loads that are considered to be non-damaging and to clip high infrequent loads that may non-conservatively bias the outcome, but care should be taken in both cases so that the test results are representative. (3)       Test duration. For any WFD-susceptible area, the average time in flight cycles and/or hours to develop WFD should first be determined. This is referred to as the WFD (average behaviour) for the subject area. The area should be modified or replaced at one third of this time unless inspection for MSD/MED is practical. If inspection is practical, that inspection should start at one third of the WFD (average behaviour), with modification/replacement at one half of that time. It is standard practice to interpret the non-factored fatigue life of one specimen as the average life. It follows that if one full‑scale fatigue test article survives a test duration of X time without an occurrence of WFD, it can be conservatively assumed that the WFD (average behaviour) of all susceptible areas is equal to X. Based on this, and assuming that the susceptible areas are impractical to inspect for MSD/MED, the replacement or modification should be implemented at X/3. For areas where MSD/MED inspections are practical, replacement/modification could be deferred until X/2, but MSD/MED inspections would have to start at X/3. The procedure should be kept in mind when deciding what the test duration will be. (4)       Post-test evaluation. One of the primary objectives of the full-scale fatigue test is to generate the data needed to determine the absolute WFD (average behaviour) for each susceptible area, or to establish a lower bound. Recall that the definition of WFD (average behaviour) is the average time required for MSD/MED to initiate and grow to the point that the static strength capability of the structure is reduced to less than the residual strength requirements of CS 25.571(b). Some work is required at the end of the test to determine the strength capability of the structure, either directly or indirectly. (i)        Residual strength tests. One acceptable way to demonstrate freedom from WFD at the end of a full-scale fatigue test is to subject the article to the required residual strength loads specified in CS 25.571(b). If the test article sustains the loads, it can be concluded that the point of WFD has yet to be reached for any of the susceptible areas. However, because fatigue cracks that might exist at the end of the test are not quantified, it is not possible to determine how far beyond the test duration WFD would occur in any of the susceptible areas without accomplishing additional work (e.g. teardown inspection). Additionally, metallic test articles may be non‑conservatively compromised relative to their future fatigue performance if static loads in excess of representative operational loads are applied. Residual strength testing could preclude the possibility of using an article for additional fatigue testing. (ii)       Teardown inspections. The residual strength capability may be evaluated indirectly by performing teardown inspections to quantify the size of any MSD/MED cracks that might be present, or to establish a lower bound on crack size based on the capability of the inspection method. Once this is done, the residual strength capability can be estimated analytically. Depending on the results, crack growth analyses may also be required to project backwards or forwards in time to estimate the WFD (average behaviour) for an area. As a minimum, teardown inspection methods should be capable of detecting the minimum size of MSD or MED cracking that would result in a WFD condition (i.e. residual strength degraded to less than the level specified in CS 25.571(b)). Ideally, it is recommended that inspection methods should be used that are capable of detecting MSD/MED cracking before it degrades the strength to less than the required level. Effective teardown inspections that are required to demonstrate freedom from WFD typically require significant resources. They typically require disassembly (e.g. fastener removal) and destruction of the test article. All areas that are or may be susceptible to WFD should be identified and examined. **(d)       Scope of full-scale fatigue test article** The following examples offer some guidance on the types of data sets that might constitute ‘sufficient evidence’ for some kinds of certification projects. The scope of the test specimen and the duration of the test are considered. (1)       New type designs Normally, this type of project would necessitate its own full-scale fatigue test of the complete airframe to represent the new structure and its loading environment. Nevertheless, prior full-scale fatigue test evidence from earlier tests performed by the applicant, or others, may also be used, and could supplement additional tests on the new model. Ultimately, the evidence needs to be sufficient to conclude with confidence that, within the LOV of the airframe, WFD will not occur. Factors 1 to 4 should be considered in determining the sufficiency of the evidence. A test duration of a minimum of twice the LOV for the aeroplane model would normally be necessary if the loading spectrum is realistic, the design and construction for the test article principal structure are the same as for the certified aeroplane, and the post-test teardown is exhaustive. If conformance to Factors 1 through 3 is less than ideal, a significantly longer test duration would be needed to conclude with confidence that WFD will not occur within the LOV. Moreover, no amount of fatigue testing will suffice if conformance to Factors 1 through 3 above is not reasonable. Consideration should also be given to the possible future need for life extension or product development, such as potential weight increases, etc. (2)       Derivative models The default position would be to test the entire airframe. However, it may be possible to reliably determine the occurrence of WFD for all or part of the derivative model from the data that the applicant generated or assembled during the original certification project. Nevertheless, the evidence needs to be sufficient to allow confidence in the calculations which show that WFD will not occur within the LOV of the aeroplane. Factors 1 through 5 should be considered in determining the sufficiency of the evidence for derivative models. For example, a change in the structural design concept, a change in the aerodynamic contours, or a modification of a structure that has a complex internal load distribution might well make analytical extrapolation from the existing full-scale fatigue test evidence very uncertain. Such changes might well necessitate full-scale fatigue testing of the actual derivative principal structure. On the other hand, a typical derivative often involves extending the fuselage by inserting ‘fuselage plugs’ that consist of a copy of the typical semi-monocoque construction for that model, with slightly modified material gauges. Normally this type of project would not necessitate its own full-scale fatigue test, particularly if very similar load paths and operating stress levels are retained. (3)       Type design changes — SBs Normally, this type of project would not necessitate its own full-scale fatigue test because the applicant would have generated, or assembled, sufficient full-scale fatigue test evidence during the original certification project that could be applied to the change. Nevertheless, as cited in the previous example, the evidence needs to be sufficient to allow confidence in the calculations which show that WFD will not occur within the LOV of the aeroplane. In addition, Factor 5, ‘The size and complexity of a design change’, should be considered. (4)       Type design changes — STCs (i)        Sufficient full-scale test evidence for structures certified under an STC may necessitate additional full-scale fatigue testing, although the extent of the design change may be small enough to use Factor 5 to establish the sufficiency of the existing full-scale fatigue test evidence. The applicant for an STC may not have access to the original equipment manufacturer (OEM)’s full-scale fatigue test data. For aeroplane types for which an LOV has been published, the STC applicant may assume that the basic structure was free from WFD up to the LOV, unless EASA has taken AD action, or intends to take action (by a proposed AD) to alleviate a WFD condition, or inspections or modifications exist in the ALS relating to WFD conditions. For the purpose of the STC applicant’s demonstration that WFD will not occur on its modification (or the underlying original structure) within the LOV, it may be assumed that the model types, to which the LOV is applicable, have received at least two full LOVs of fatigue testing, under realistic loads, and have received thorough post-test inspections that did not detect any WFD, or the ALS includes from the outset details of the modifications required to address WFD that will need specific consideration by the STC applicant. With this knowledge, and Factors 1 through 5, the STC applicant may be able to demonstrate that WFD will not occur on its modification (or the underlying original structure) within the LOV. If, however, the modification significantly affects the distribution of stress in the underlying structure, or significantly alters loads in other parts of the aeroplane, or significantly alters the intended mission of the aeroplane, or if the modification is significantly different in its structural concept from the certified aeroplane being modified, additional representative fatigue test evidence would be necessary. (ii)       In addition, Factor 6 ‘The age of the aeroplane being modified’ comes into play for modifications made to older aeroplanes. The STC applicant should demonstrate freedom from WFD up to the LOV of the aeroplane being modified. For example, an applicant for an STC to an aeroplane that has reached an age equivalent to 75 % of its LOV should demonstrate that the modified aeroplane will be free from WFD for at least the remaining 25 % of the LOV. Although an applicant could attempt to demonstrate freedom from WFD for a longer period, this may not be possible unless the OEM cooperates by providing data for the basic structure. A short DSG for the modification could simplify the demonstration of freedom from WFD for the STC applicant. Nevertheless, the applicant should also be aware that the LOV of the aeroplane is not a fixed life; it may be extended as a result of a structural re‑evaluation and service action plan, such as those developed for certain models under the FAA’s ‘Aging Aircraft Program’. Unless the modifier also re-evaluates its STC modification, the shorter goal for the modification could impede extending the LOV of the modified aeroplanes. (5)       Major repairs. New repairs (for which the applicable certification basis requires WFD evaluation) that differ from the repairs contained in the OEM’s SRM, but that are equivalent in design from such repairs, and that meet CS-25 specifications in other respects, would not necessitate full-scale fatigue testing to support freedom from WFD up to the LOV. Major repair solutions (that may be susceptible to WFD) which utilise design concepts (e.g. new materials, other production processes, new design details) different from the previously approved repair data may need further testing. **(e)       Use of existing full-scale fatigue test data** In some cases, especially for establishing an LOV in accordance with point 26.303 of Part-26, or for derivative models and type design changes accomplished by the TCH, there may be existing full-scale fatigue test data that may be used to support compliance and mitigate the need to perform additional testing. Any physical differences between the structure originally tested and the structure being considered that could affect its fatigue behaviour must be identified and reconciled. Differences that should be addressed include, but are not limited to, differences in any of the physical attributes listed under point (c)(1) of this Annex and differences in operational loading. Typical developments that affect the applicability of the original LOV demonstration data are the: —        gross weight (e.g. if it increases), —        cabin pressurisation (e.g. a change in the maximum cabin or operating altitude), or —        flight segment parameters. The older the test data, the harder it may be to demonstrate that it is sufficient. Often test articles were not conformed, neither were test plans or reports submitted to EASA as part of the compliance data package. The rigour of loading sequences has varied significantly over the years, and from OEM to OEM. Additionally, testing philosophies and protocols were not standardised. For example, post-test evaluations, if any, varied significantly and in some cases consisted of nothing more than limited visual inspections. However, there may be acceptable data from the early full-scale fatigue tests that the applicant proposes to use to support compliance. In order to use such data, the configuration of the test article and the loading must be verified, and the issue of the residual strength capability of the article (or teardown data) at the end of the test must be addressed. **(f)        Use of in-service data** There may be in-service data that can be used to support WFD evaluations. Examples of such data are as follows: —        Documented positive findings of MSD/MED cracks that include the location, size and the time in service of the affected aeroplane, along with a credible record of how the aircraft had been operated since the original delivery. —        Documented negative findings from in-service inspections for MSD/MED cracks on a statistically significant number of aeroplanes, with the time in service of each aircraft, and a credible record of how each aircraft had been operated since the original delivery. For this data to be useful, the inspection methods used should have been capable of detecting MSD/MED crack sizes equal to or smaller than those sizes that could reduce the strength of the structure to less than the residual strength levels specified in CS 25.571(b). —        Documented findings from the destructive teardown inspection of structures from in‑service aircraft. This might be structures (e.g. fuselage splices) removed from the aircraft that were subsequently returned to service, or from retired aircraft. It would also be necessary to have a credible record of the operational loading experienced by the subject structure up to the time it was taken out of service. Prior to using in-service data, any physical or loading differences that exist between the structure of the in-service or retired aeroplanes and the structure being certified should be identified and reconciled as discussed above. [Amdt 20/20] [Amdt 20/22] Annex 1 to Appendix 2 to AMC 20-20B Full-scale fatigue test evidence ED Decision 2021/007/R Contents (a) Overview (b) Full-scale fatigue test evidence (c) Key elements of a full-scale fatigue test programme (1) Article (2) Test set-up and loading (i) Test set-up (ii) Test loading (3) Test duration (i) New type certificates and derivatives (ii) Repairs and type design changes (4) Post-test evaluation (i) Residual strength tests (ii) Teardown inspections (d) Scope of full-scale fatigue test article (1) New type designs (2) Derivative models (3) Type design changes — SBs (4) Type design changes — STCs (5) Major repairs (e) Use of existing full-scale fatigue test data (f) Use of in-service data (a) Overview CS 25.571(b) Amendment 19 specifies that special consideration for WFD must be included in the fatigue and damage tolerance evaluation where the design is such that this type of damage could occur. CS 25.571(b) Amendment 19 specifies the effectiveness of the provisions to preclude the possibility of WFD occurring within the limits of validity of the maintenance programme to be demonstrated with sufficient full-scale fatigue test evidence. The determination of what constitutes ‘sufficient full-scale test evidence’ requires a considerable amount of engineering judgement and is a matter that should be discussed and agreed to between an applicant and EASA early in the planning stage of a certification project. Sufficient test evidence is also necessary to support compliance with CS 26.303 and the most straightforward means of compliance is to utilise existing full-scale test evidence. (b) Full-scale fatigue test evidence In general, sufficient full-scale fatigue test evidence consists of full-scale fatigue testing to at least twice the LOV, followed by specific inspections and analyses to determine that WFD has not occurred. The following factors should be considered in determining the sufficiency of the evidence: Factor 1: The comparability of the load spectrum between the test and the projected usage of the aeroplane. Factor 2: The comparability of the airframe materials, design and build standards between the test article and the certified aeroplane. Factor 3: The extent of post-test teardown inspection, residual strength testing and analysis for determining whether widespread fatigue cracking has occurred. Factor 4: The duration of the fatigue testing. Factor 5: The size and complexity of a design or build standard change. This factor applies to design changes made to a model that has already been certified and for which full-scale fatigue test evidence for the original structure should have already been determined to be sufficient. Small, simple design changes, comparable to the original structure, could be analytically determined to be equivalent to the original structure in their propensity for WFD. In such cases, additional full-scale fatigue test evidence should not be necessary. Factor 6: In the case of major changes and STCs, the age of an aeroplane being modified. This factor applies to aeroplanes that have already accumulated a portion of their LOV prior to being modified. An applicant should demonstrate freedom from WFD up to the LOV in place for the original aeroplane and may take into account the age of the aeroplane being modified. (c) Key elements of a full-scale fatigue test programme The following guidance addresses key elements of a test programme that is intended to generate the data necessary to support compliance, and it can also be used to evaluate and interpret existing full-scale test data for the purposes of supporting compliance with point 26.303 of Part-26. (1) Article. The test article should be representative of the structure of the aeroplane to be evaluated (i.e. ideally a production-standard article). The attributes of the type design that could affect MSD/MED initiation, growth and subsequent residual strength capability should be replicated as closely as possible on the test article. Critical attributes include, but are not limited to, the following: — material types and forms; — dimensions; — joining methods and details; — coating and plating; — the use of faying surface sealant; — assembly processes and sequences; and — the influence of secondary structure (e.g. loads induced due to proximity to the structure under evaluation). (2) Test set-up and loading. The test set-up and loading should result in a realistic simulation of the expected operational loads. (i) Test set-up. The test set-up dictates how loads are introduced into the structure and reacted. Every effort should be made to introduce and react loads as realistically as possible. When a compromise is made (e.g. wing air loading), the resulting internal loads should be evaluated (e.g. using finite element methods) to ensure that the structure is not being unrealistically underloaded or overloaded, locally or globally. (ii) Loading spectrum. The test loading spectrum should include loads from all damaging sources (e.g. cabin pressurisation, manoeuvres, gusts, engine thrust, control surface deflections, and landing impacts) that are significant for the structure being evaluated. Consideration should also be given to temperature and other environmental effects that may affect internal loads. A supporting rationale should be provided when a load source is not represented in a sequence. Additionally, differences between the test sequence and the expected operational sequence should be justified. For example, it is standard practice to eliminate low loads that are considered to be non-damaging and to clip high infrequent loads that may non-conservatively bias the outcome, but care should be taken in both cases so that the test results are representative. (3) Test duration. For any WFD-susceptible area, the average time in flight cycles and/or hours to develop WFD should first be determined. This is referred to as the WFD (average behaviour) for the subject area. The area should be modified or replaced at one third of this time unless inspection for MSD/MED is practical. If inspection is practical, that inspection should start at one third of the WFD (average behaviour), with modification/replacement at one half of that time. It is standard practice to interpret the non-factored fatigue life of one specimen as the average life. It follows that if one full‑scale fatigue test article survives a test duration of X time without an occurrence of WFD, it can be conservatively assumed that the WFD (average behaviour) of all susceptible areas is equal to X. Based on this, and assuming that the susceptible areas are impractical to inspect for MSD/MED, the replacement or modification should be implemented at X/3. For areas where MSD/MED inspections are practical, replacement/modification could be deferred until X/2, but MSD/MED inspections would have to start at X/3. The procedure should be kept in mind when deciding what the test duration will be. (4) Post-test evaluation. One of the primary objectives of the full-scale fatigue test is to generate the data needed to determine the absolute WFD (average behaviour) for each susceptible area, or to establish a lower bound. Recall that the definition of WFD (average behaviour) is the average time required for MSD/MED to initiate and grow to the point that the static strength capability of the structure is reduced to less than the residual strength requirements of CS 25.571(b). Some work is required at the end of the test to determine the strength capability of the structure, either directly or indirectly. (i) Residual strength tests. One acceptable way to demonstrate freedom from WFD at the end of a full-scale fatigue test is to subject the article to the required residual strength loads specified in CS 25.571(b). If the test article sustains the loads, it can be concluded that the point of WFD has yet to be reached for any of the susceptible areas. However, because fatigue cracks that might exist at the end of the test are not quantified, it is not possible to determine how far beyond the test duration WFD would occur in any of the susceptible areas without accomplishing additional work (e.g. teardown inspection). Additionally, metallic test articles may be non‑conservatively compromised relative to their future fatigue performance if static loads in excess of representative operational loads are applied. Residual strength testing could preclude the possibility of using an article for additional fatigue testing. (ii) Teardown inspections. The residual strength capability may be evaluated indirectly by performing teardown inspections to quantify the size of any MSD/MED cracks that might be present, or to establish a lower bound on crack size based on the capability of the inspection method. Once this is done, the residual strength capability can be estimated analytically. Depending on the results, crack growth analyses may also be required to project backwards or forwards in time to estimate the WFD (average behaviour) for an area. As a minimum, teardown inspection methods should be capable of detecting the minimum size of MSD or MED cracking that would result in a WFD condition (i.e. residual strength degraded to less than the level specified in CS 25.571(b)). Ideally, it is recommended that inspection methods should be used that are capable of detecting MSD/MED cracking before it degrades the strength to less than the required level. Effective teardown inspections that are required to demonstrate freedom from WFD typically require significant resources. They typically require disassembly (e.g. fastener removal) and destruction of the test article. All areas that are or may be susceptible to WFD should be identified and examined. **(d) Scope of full-scale fatigue test article** The following examples offer some guidance on the types of data sets that might constitute ‘sufficient evidence’ for some kinds of certification projects. The scope of the test specimen and the duration of the test are considered. (1) New type designs Normally, this type of project would necessitate its own full-scale fatigue test of the complete airframe to represent the new structure and its loading environment. Nevertheless, prior full-scale fatigue test evidence from earlier tests performed by the applicant, or others, may also be used, and could supplement additional tests on the new model. Ultimately, the evidence needs to be sufficient to conclude with confidence that, within the LOV of the airframe, WFD will not occur. Factors 1 to 4 should be considered in determining the sufficiency of the evidence. A test duration of a minimum of twice the LOV for the aeroplane model would normally be necessary if the loading spectrum is realistic, the design and construction for the test article principal structure are the same as for the certified aeroplane, and the post-test teardown is exhaustive. If conformance to Factors 1 through 3 is less than ideal, a significantly longer test duration would be needed to conclude with confidence that WFD will not occur within the LOV. Moreover, no amount of fatigue testing will suffice if conformance to Factors 1 through 3 above is not reasonable. Consideration should also be given to the possible future need for life extension or product development, such as potential weight increases, etc. (2) Derivative models The default position would be to test the entire airframe. However, it may be possible to reliably determine the occurrence of WFD for all or part of the derivative model from the data that the applicant generated or assembled during the original certification project. Nevertheless, the evidence needs to be sufficient to allow confidence in the calculations which show that WFD will not occur within the LOV of the aeroplane. Factors 1 through 5 should be considered in determining the sufficiency of the evidence for derivative models. For example, a change in the structural design concept, a change in the aerodynamic contours, or a modification of a structure that has a complex internal load distribution might well make analytical extrapolation from the existing full-scale fatigue test evidence very uncertain. Such changes might well necessitate full-scale fatigue testing of the actual derivative principal structure. On the other hand, a typical derivative often involves extending the fuselage by inserting ‘fuselage plugs’ that consist of a copy of the typical semi-monocoque construction for that model, with slightly modified material gauges. Normally this type of project would not necessitate its own full-scale fatigue test, particularly if very similar load paths and operating stress levels are retained. (3) Type design changes — SBs Normally, this type of project would not necessitate its own full-scale fatigue test because the applicant would have generated, or assembled, sufficient full-scale fatigue test evidence during the original certification project that could be applied to the change. Nevertheless, as cited in the previous example, the evidence needs to be sufficient to allow confidence in the calculations which show that WFD will not occur within the LOV of the aeroplane. In addition, Factor 5, ‘The size and complexity of a design change’, should be considered. (4) Type design changes — STCs (i) Sufficient full-scale test evidence for structures certified under an STC may necessitate additional full-scale fatigue testing, although the extent of the design change may be small enough to use Factor 5 to establish the sufficiency of the existing full-scale fatigue test evidence. The applicant for an STC may not have access to the original equipment manufacturer (OEM)’s full-scale fatigue test data. For aeroplane types for which an LOV has been published, the STC applicant may assume that the basic structure was free from WFD up to the LOV, unless EASA has taken AD action, or intends to take action (by a proposed AD) to alleviate a WFD condition, or inspections or modifications exist in the ALS relating to WFD conditions. For the purpose of the STC applicant’s demonstration that WFD will not occur on its modification (or the underlying original structure) within the LOV, it may be assumed that the model types, to which the LOV is applicable, have received at least two full LOVs of fatigue testing, under realistic loads, and have received thorough post-test inspections that did not detect any WFD, or the ALS includes from the outset details of the modifications required to address WFD that will need specific consideration by the STC applicant. With this knowledge, and Factors 1 through 5, the STC applicant may be able to demonstrate that WFD will not occur on its modification (or the underlying original structure) within the LOV. If, however, the modification significantly affects the distribution of stress in the underlying structure, or significantly alters loads in other parts of the aeroplane, or significantly alters the intended mission of the aeroplane, or if the modification is significantly different in its structural concept from the certified aeroplane being modified, additional representative fatigue test evidence would be necessary. (ii) In addition, Factor 6 ‘The age of the aeroplane being modified’ comes into play for modifications made to older aeroplanes. The STC applicant should demonstrate freedom from WFD up to the LOV of the aeroplane being modified. For example, an applicant for an STC to an aeroplane that has reached an age equivalent to 75 % of its LOV should demonstrate that the modified aeroplane will be free from WFD for at least the remaining 25 % of the LOV. Although an applicant could attempt to demonstrate freedom from WFD for a longer period, this may not be possible unless the OEM cooperates by providing data for the basic structure. A short DSG for the modification could simplify the demonstration of freedom from WFD for the STC applicant. Nevertheless, the applicant should also be aware that the LOV of the aeroplane is not a fixed life; it may be extended as a result of a structural re‑evaluation and service action plan, such as those developed for certain models under the FAA’s ‘Aging Aircraft Program’. Unless the modifier also re-evaluates its STC modification, the shorter goal for the modification could impede extending the LOV of the modified aeroplanes. (5) Major repairs. New repairs (for which the applicable certification basis requires WFD evaluation) that differ from the repairs contained in the OEM’s SRM, but that are equivalent in design from such repairs, and that meet CS-25 specifications in other respects, would not necessitate full-scale fatigue testing to support freedom from WFD up to the LOV. Major repair solutions (that may be susceptible to WFD) which utilise design concepts (e.g. new materials, other production processes, new design details) different from the previously approved repair data may need further testing. **(e) Use of existing full-scale fatigue test data** In some cases, especially for establishing an LOV in accordance with point 26.303 of Part-26, or for derivative models and type design changes accomplished by the TCH, there may be existing full-scale fatigue test data that may be used to support compliance and mitigate the need to perform additional testing. Any physical differences between the structure originally tested and the structure being considered that could affect its fatigue behaviour must be identified and reconciled. Differences that should be addressed include, but are not limited to, differences in any of the physical attributes listed under point (c)(1) of this Annex and differences in operational loading. Typical developments that affect the applicability of the original LOV demonstration data are the: — gross weight (e.g. if it increases), — cabin pressurisation (e.g. a change in the maximum cabin or operating altitude), or — flight segment parameters. The older the test data, the harder it may be to demonstrate that it is sufficient. Often test articles were not conformed, neither were test plans or reports submitted to EASA as part of the compliance data package. The rigour of loading sequences has varied significantly over the years, and from OEM to OEM. Additionally, testing philosophies and protocols were not standardised. For example, post-test evaluations, if any, varied significantly and in some cases consisted of nothing more than limited visual inspections. However, there may be acceptable data from the early full-scale fatigue tests that the applicant proposes to use to support compliance. In order to use such data, the configuration of the test article and the loading must be verified, and the issue of the residual strength capability of the article (or teardown data) at the end of the test must be addressed. **(f) Use of in-service data** There may be in-service data that can be used to support WFD evaluations. Examples of such data are as follows: — Documented positive findings of MSD/MED cracks that include the location, size and the time in service of the affected aeroplane, along with a credible record of how the aircraft had been operated since the original delivery. — Documented negative findings from in-service inspections for MSD/MED cracks on a statistically significant number of aeroplanes, with the time in service of each aircraft, and a credible record of how each aircraft had been operated since the original delivery. For this data to be useful, the inspection methods used should have been capable of detecting MSD/MED crack sizes equal to or smaller than those sizes that could reduce the strength of the structure to less than the residual strength levels specified in CS 25.571(b). — Documented findings from the destructive teardown inspection of structures from in‑service aircraft. This might be structures (e.g. fuselage splices) removed from the aircraft that were subsequently returned to service, or from retired aircraft. It would also be necessary to have a credible record of the operational loading experienced by the subject structure up to the time it was taken out of service. Prior to using in-service data, any physical or loading differences that exist between the structure of the in-service or retired aeroplanes and the structure being certified should be identified and reconciled as discussed above. [Amdt 20/20] [Amdt 20/22] Annex 1 to Appendix 2 to AMC 20-20A Full-scale fatigue test evidence ED Decision 2020/023/R Contents (a)       Overview (b)       Full-scale fatigue test evidence (c)       Key elements of a full-scale fatigue test programme (1)       Article (2)       Test set-up and loading (i)        Test set-up (ii)       Test loading (3)       Test duration (i)        New type certificates and derivatives (ii)       Repairs and type design changes (4)       Post-test evaluation (i)        Residual strength tests (ii)       Teardown inspections (d)       Scope of full-scale fatigue test article (1)       New type designs (2)       Derivative models (3)       Type design changes — SBs (4)       Type design changes — STCs (5)       Major repairs (e)       Use of existing full-scale fatigue test data (f)        Use of in-service data (a)       Overview CS 25.571(b) Amendment 19 specifies that special consideration for WFD must be included in the fatigue and damage tolerance evaluation where the design is such that this type of damage could occur. CS 25.571(b) Amendment 19 specifies the effectiveness of the provisions to preclude the possibility of WFD occurring within the limits of validity of the maintenance programme to be demonstrated with sufficient full-scale fatigue test evidence. The determination of what constitutes ‘sufficient full-scale test evidence’ requires a considerable amount of engineering judgement and is a matter that should be discussed and agreed to between an applicant and EASA early in the planning stage of a certification project. Sufficient test evidence is also necessary to support compliance with CS 26.303 and the most straightforward means of compliance is to utilise existing full-scale test evidence. (b)       Full-scale fatigue test evidence In general, sufficient full-scale fatigue test evidence consists of full-scale fatigue testing to at least twice the LOV, followed by specific inspections and analyses to determine that WFD has not occurred. The following factors should be considered in determining the sufficiency of the evidence: Factor 1: The comparability of the load spectrum between the test and the projected usage of the aeroplane. Factor 2: The comparability of the airframe materials, design and build standards between the test article and the certified aeroplane. Factor 3: The extent of post-test teardown inspection, residual strength testing and analysis for determining whether widespread fatigue cracking has occurred. Factor 4: The duration of the fatigue testing. Factor 5: The size and complexity of a design or build standard change. This factor applies to design changes made to a model that has already been certified and for which full-scale fatigue test evidence for the original structure should have already been determined to be sufficient. Small, simple design changes, comparable to the original structure, could be analytically determined to be equivalent to the original structure in their propensity for WFD. In such cases, additional full-scale fatigue test evidence should not be necessary. Factor 6: In the case of major changes and STCs, the age of an aeroplane being modified. This factor applies to aeroplanes that have already accumulated a portion of their LOV prior to being modified. An applicant should demonstrate freedom from WFD up to the LOV in place for the original aeroplane and may take into account the age of the aeroplane being modified. (c)       Key elements of a full-scale fatigue test programme The following guidance addresses key elements of a test programme that is intended to generate the data necessary to support compliance, and it can also be used to evaluate and interpret existing full-scale test data for the purposes of supporting compliance with point 26.303 of Part-26. (1)       Article. The test article should be representative of the structure of the aeroplane to be evaluated (i.e. ideally a production-standard article). The attributes of the type design that could affect MSD/MED initiation, growth and subsequent residual strength capability should be replicated as closely as possible on the test article. Critical attributes include, but are not limited to, the following: —        material types and forms; —        dimensions; —        joining methods and details; —        coating and plating; —        the use of faying surface sealant; —        assembly processes and sequences; and —        the influence of secondary structure (e.g. loads induced due to proximity to the structure under evaluation). (2)       Test set-up and loading. The test set-up and loading should result in a realistic simulation of the expected operational loads. (i)        Test set-up. The test set-up dictates how loads are introduced into the structure and reacted. Every effort should be made to introduce and react loads as realistically as possible. When a compromise is made (e.g. wing air loading), the resulting internal loads should be evaluated (e.g. using finite element methods) to ensure that the structure is not being unrealistically underloaded or overloaded, locally or globally. (ii)       Loading spectrum. The test loading spectrum should include loads from all damaging sources (e.g. cabin pressurisation, manoeuvres, gusts, engine thrust, control surface deflections, and landing impacts) that are significant for the structure being evaluated. Consideration should also be given to temperature and other environmental effects that may affect internal loads. A supporting rationale should be provided when a load source is not represented in a sequence. Additionally, differences between the test sequence and the expected operational sequence should be justified. For example, it is standard practice to eliminate low loads that are considered to be non-damaging and to clip high infrequent loads that may non-conservatively bias the outcome, but care should be taken in both cases so that the test results are representative. (3)       Test duration. For any WFD-susceptible area, the average time in flight cycles and/or hours to develop WFD should first be determined. This is referred to as the WFD (average behaviour) for the subject area. The area should be modified or replaced at one third of this time unless inspection for MSD/MED is practical. If inspection is practical, that inspection should start at one third of the WFD (average behaviour), with modification/replacement at one half of that time. It is standard practice to interpret the non-factored fatigue life of one specimen as the average life. It follows that if one full‑scale fatigue test article survives a test duration of X time without an occurrence of WFD, it can be conservatively assumed that the WFD (average behaviour) of all susceptible areas is equal to X. Based on this, and assuming that the susceptible areas are impractical to inspect for MSD/MED, the replacement or modification should be implemented at X/3. For areas where MSD/MED inspections are practical, replacement/modification could be deferred until X/2, but MSD/MED inspections would have to start at X/3. The procedure should be kept in mind when deciding what the test duration will be. (4)       Post-test evaluation. One of the primary objectives of the full-scale fatigue test is to generate the data needed to determine the absolute WFD (average behaviour) for each susceptible area, or to establish a lower bound. Recall that the definition of WFD (average behaviour) is the average time required for MSD/MED to initiate and grow to the point that the static strength capability of the structure is reduced to less than the residual strength requirements of CS 25.571(b). Some work is required at the end of the test to determine the strength capability of the structure, either directly or indirectly. (i)        Residual strength tests. One acceptable way to demonstrate freedom from WFD at the end of a full-scale fatigue test is to subject the article to the required residual strength loads specified in CS 25.571(b). If the test article sustains the loads, it can be concluded that the point of WFD has yet to be reached for any of the susceptible areas. However, because fatigue cracks that might exist at the end of the test are not quantified, it is not possible to determine how far beyond the test duration WFD would occur in any of the susceptible areas without accomplishing additional work (e.g. teardown inspection). Additionally, metallic test articles may be non‑conservatively compromised relative to their future fatigue performance if static loads in excess of representative operational loads are applied. Residual strength testing could preclude the possibility of using an article for additional fatigue testing. (ii)       Teardown inspections. The residual strength capability may be evaluated indirectly by performing teardown inspections to quantify the size of any MSD/MED cracks that might be present, or to establish a lower bound on crack size based on the capability of the inspection method. Once this is done, the residual strength capability can be estimated analytically. Depending on the results, crack growth analyses may also be required to project backwards or forwards in time to estimate the WFD (average behaviour) for an area. As a minimum, teardown inspection methods should be capable of detecting the minimum size of MSD or MED cracking that would result in a WFD condition (i.e. residual strength degraded to less than the level specified in CS 25.571(b)). Ideally, it is recommended that inspection methods should be used that are capable of detecting MSD/MED cracking before it degrades the strength to less than the required level. Effective teardown inspections that are required to demonstrate freedom from WFD typically require significant resources. They typically require disassembly (e.g. fastener removal) and destruction of the test article. All areas that are or may be susceptible to WFD should be identified and examined. **(d)       Scope of full-scale fatigue test article** The following examples offer some guidance on the types of data sets that might constitute ‘sufficient evidence’ for some kinds of certification projects. The scope of the test specimen and the duration of the test are considered. (1)       New type designs Normally, this type of project would necessitate its own full-scale fatigue test of the complete airframe to represent the new structure and its loading environment. Nevertheless, prior full-scale fatigue test evidence from earlier tests performed by the applicant, or others, may also be used, and could supplement additional tests on the new model. Ultimately, the evidence needs to be sufficient to conclude with confidence that, within the LOV of the airframe, WFD will not occur. Factors 1 to 4 should be considered in determining the sufficiency of the evidence. A test duration of a minimum of twice the LOV for the aeroplane model would normally be necessary if the loading spectrum is realistic, the design and construction for the test article principal structure are the same as for the certified aeroplane, and the post-test teardown is exhaustive. If conformance to Factors 1 through 3 is less than ideal, a significantly longer test duration would be needed to conclude with confidence that WFD will not occur within the LOV. Moreover, no amount of fatigue testing will suffice if conformance to Factors 1 through 3 above is not reasonable. Consideration should also be given to the possible future need for life extension or product development, such as potential weight increases, etc. (2)       Derivative models The default position would be to test the entire airframe. However, it may be possible to reliably determine the occurrence of WFD for all or part of the derivative model from the data that the applicant generated or assembled during the original certification project. Nevertheless, the evidence needs to be sufficient to allow confidence in the calculations which show that WFD will not occur within the LOV of the aeroplane. Factors 1 through 5 should be considered in determining the sufficiency of the evidence for derivative models. For example, a change in the structural design concept, a change in the aerodynamic contours, or a modification of a structure that has a complex internal load distribution might well make analytical extrapolation from the existing full-scale fatigue test evidence very uncertain. Such changes might well necessitate full-scale fatigue testing of the actual derivative principal structure. On the other hand, a typical derivative often involves extending the fuselage by inserting ‘fuselage plugs’ that consist of a copy of the typical semi-monocoque construction for that model, with slightly modified material gauges. Normally this type of project would not necessitate its own full-scale fatigue test, particularly if very similar load paths and operating stress levels are retained. (3)       Type design changes — SBs Normally, this type of project would not necessitate its own full-scale fatigue test because the applicant would have generated, or assembled, sufficient full-scale fatigue test evidence during the original certification project that could be applied to the change. Nevertheless, as cited in the previous example, the evidence needs to be sufficient to allow confidence in the calculations which show that WFD will not occur within the LOV of the aeroplane. In addition, Factor 5, ‘The size and complexity of a design change’, should be considered. (4)       Type design changes — STCs (i)        Sufficient full-scale test evidence for structures certified under an STC may necessitate additional full-scale fatigue testing, although the extent of the design change may be small enough to use Factor 5 to establish the sufficiency of the existing full-scale fatigue test evidence. The applicant for an STC may not have access to the original equipment manufacturer (OEM)’s full-scale fatigue test data. For aeroplane types for which an LOV has been published, the STC applicant may assume that the basic structure was free from WFD up to the LOV, unless EASA has taken AD action, or intends to take action (by a proposed AD) to alleviate a WFD condition, or inspections or modifications exist in the ALS relating to WFD conditions. For the purpose of the STC applicant’s demonstration that WFD will not occur on its modification (or the underlying original structure) within the LOV, it may be assumed that the model types, to which the LOV is applicable, have received at least two full LOVs of fatigue testing, under realistic loads, and have received thorough post-test inspections that did not detect any WFD, or the ALS includes from the outset details of the modifications required to address WFD that will need specific consideration by the STC applicant. With this knowledge, and Factors 1 through 5, the STC applicant may be able to demonstrate that WFD will not occur on its modification (or the underlying original structure) within the LOV. If, however, the modification significantly affects the distribution of stress in the underlying structure, or significantly alters loads in other parts of the aeroplane, or significantly alters the intended mission of the aeroplane, or if the modification is significantly different in its structural concept from the certified aeroplane being modified, additional representative fatigue test evidence would be necessary. (ii)       In addition, Factor 6 ‘The age of the aeroplane being modified’ comes into play for modifications made to older aeroplanes. The STC applicant should demonstrate freedom from WFD up to the LOV of the aeroplane being modified. For example, an applicant for an STC to an aeroplane that has reached an age equivalent to 75 % of its LOV should demonstrate that the modified aeroplane will be free from WFD for at least the remaining 25 % of the LOV. Although an applicant could attempt to demonstrate freedom from WFD for a longer period, this may not be possible unless the OEM cooperates by providing data for the basic structure. A short DSG for the modification could simplify the demonstration of freedom from WFD for the STC applicant. Nevertheless, the applicant should also be aware that the LOV of the aeroplane is not a fixed life; it may be extended as a result of a structural re‑evaluation and service action plan, such as those developed for certain models under the FAA’s ‘Aging Aircraft Program’. Unless the modifier also re-evaluates its STC modification, the shorter goal for the modification could impede extending the LOV of the modified aeroplanes. (5)       Major repairs. New repairs (for which the applicable certification basis requires WFD evaluation) that differ from the repairs contained in the OEM’s SRM, but that are equivalent in design from such repairs, and that meet CS-25 specifications in other respects, would not necessitate full-scale fatigue testing to support freedom from WFD up to the LOV. Major repair solutions (that may be susceptible to WFD) which utilise design concepts (e.g. new materials, other production processes, new design details) different from the previously approved repair data may need further testing. **(e)       Use of existing full-scale fatigue test data** In some cases, especially for establishing an LOV in accordance with point 26.303 of Part-26, or for derivative models and type design changes accomplished by the TCH, there may be existing full-scale fatigue test data that may be used to support compliance and mitigate the need to perform additional testing. Any physical differences between the structure originally tested and the structure being considered that could affect its fatigue behaviour must be identified and reconciled. Differences that should be addressed include, but are not limited to, differences in any of the physical attributes listed under point (c)(1) of this Annex and differences in operational loading. Typical developments that affect the applicability of the original LOV demonstration data are the: —        gross weight (e.g. if it increases), —        cabin pressurisation (e.g. a change in the maximum cabin or operating altitude), or —        flight segment parameters. The older the test data, the harder it may be to demonstrate that it is sufficient. Often test articles were not conformed, neither were test plans or reports submitted to EASA as part of the compliance data package. The rigour of loading sequences has varied significantly over the years, and from OEM to OEM. Additionally, testing philosophies and protocols were not standardised. For example, post-test evaluations, if any, varied significantly and in some cases consisted of nothing more than limited visual inspections. However, there may be acceptable data from the early full-scale fatigue tests that the applicant proposes to use to support compliance. In order to use such data, the configuration of the test article and the loading must be verified, and the issue of the residual strength capability of the article (or teardown data) at the end of the test must be addressed. **(f)        Use of in-service data** There may be in-service data that can be used to support WFD evaluations. Examples of such data are as follows: —        Documented positive findings of MSD/MED cracks that include the location, size and the time in service of the affected aeroplane, along with a credible record of how the aircraft had been operated since the original delivery. —        Documented negative findings from in-service inspections for MSD/MED cracks on a statistically significant number of aeroplanes, with the time in service of each aircraft, and a credible record of how each aircraft had been operated since the original delivery. For this data to be useful, the inspection methods used should have been capable of detecting MSD/MED crack sizes equal to or smaller than those sizes that could reduce the strength of the structure to less than the residual strength levels specified in CS 25.571(b). —        Documented findings from the destructive teardown inspection of structures from in‑service aircraft. This might be structures (e.g. fuselage splices) removed from the aircraft that were subsequently returned to service, or from retired aircraft. It would also be necessary to have a credible record of the operational loading experienced by the subject structure up to the time it was taken out of service. Prior to using in-service data, any physical or loading differences that exist between the structure of the in-service or retired aeroplanes and the structure being certified should be identified and reconciled as discussed above. [Amdt 20/20]