AMC 20-20A Continuing structural integrity programme

1.         PURPOSE

(a)       This acceptable means of compliance (AMC) provides guidance to typecertificate holders(TCHs), supplemental type certificate holders (STCHs), repair approval holders, maintenance organisations, operators and competent authorities for developing continuing structural integrity programmes to ensure safe operation of ageing aircraft throughout their operational lives.

This AMC is primarily aimed at large aeroplanes; however, this material is also applicable to other aircraft types for operators and TCHs wishing to develop robust continuing structural integrity programmes.

(b)       It is particularly important for the TCHs of ageing aircraft to ensure that their continuing structural integrity programmes remain valid throughout the operational life of the aircraft.

(c)       The means of compliance described in this document provides guidance to supplement the engineering and operational judgement that must form the basis of any compliance findings relative to continuing structural integrity programmes.

(d)       Like all acceptable means of compliance material, this AMC is not in itself mandatory, and does not constitute a requirement. It describes an acceptable means, but not the only means, for showing compliance with the requirements. While these guidelines are not mandatory, they are derived from extensive industry experience in determining compliance with the relevant requirements.

(e)       This AMC also supports compliance with the ageing structural integrity requirements in Annex I (Part-26) to Regulation (EU) 2015/640, as introduced by Regulation (EU) 2020/1159 (ref. points 26.300 through 26.370 and the associated CS-26 paragraphs) including limits of validity (LOVs), WFD evaluation, damage tolerance for repairs and modifications, and processes for ensuring the continued validity of the continuing structural integrity programme.

2.         RELATED REGULATIONS AND DOCUMENTS

(a)       Implementing Rules and Certification Specifications:

Point 21.A.61           Instructions for continued airworthiness

Point 21.A.120         Instructions for continued airworthiness

Points 26.300 through 26.334 applicable to DAHs

Point 21.A.433         Repair design

Point 26.370                          Rules applicable to operators

Point M.A.302         Maintenance programme

CS 25.571                   Damagetolerance and fatigue evaluation of structure

CS 25.1529                 Instructions for continued airworthiness

CS 26.300 through 26.370 Means of compliance for Part-26 ageing aeroplane structures requirements

(b)       EASA AMC and FAA Advisory Circulars

AMC 25.571              Damage tolerance and fatigue evaluation of structure

AC 91-81                                 Management Programs for Airplanes with Demonstrated Risk of Catastrophic Failure Due to Fatigue, 29 April 2008, FAA

AC 91-56B                  Continuing Structural Integrity for Airplanes,7 March 2008, FAA

AC 120-73                               Damage Tolerance Assessment of Repairs to Pressurised Fuselages, FAA,14 December 2000

AC 120-93                  Damage Tolerance Inspections for Repairs and Alterations

AC 120-104               Establishing and Implementing Limit of Validity to Prevent Widespread Fatigue Damage

AC 25.1529-1A         Instructions for Continued Airworthiness of StructuralRepairs on Transport Airplanes, FAA, 20 November 2007

(c)       Related documents

—      ‘Recommendations for Regulatory Action to Prevent Widespread Fatigue Damage in the Commercial Aeroplane Fleet’, Revision A, dated June 29, 1999 [A report of the Airworthiness Assurance Working Group for the Aviation Rulemaking Advisory Committee Transport Aircraft and Engine Issues].

—      AAWG Final Report on Continued Airworthiness of Structural Repairs, Dec 1996.

—      ATA report 51-93-01 structural maintenance programme guidelines for continuing airworthiness, May 1993.

—      AAWG Report on Structures Task Group Guidelines, Rev 1 June 1996.

—      AAWG Report: Recommendations concerning ARAC taskings FR Doc04-10816 Ref.: Aging Airplane safety final rule. 14 CFR 121.370a and 129.16.

—      Federal Aviation Administration 14 CFR Parts 26, 121, and 129 [Docket No FAA‑2005-21693; Amendment Nos 26–1, 121–337, 129–44] Damage Tolerance Data for Repairs and Alterations Final Rule.

—      Federal Aviation Administration 14 CFR Parts 25, 26, 121, and 129 [Docket No FAA‑2006-24281; Amendment Nos 25–132, 26–5, 121–351, 129–48] Aging Airplane Program: Widespread Fatigue Damage Final Rule.

3.         BACKGROUND

Service experience has shown there is a need to have continually updated knowledge on the structural integrity of aircraft, especially as they become older,to ensure they continue to meet the level of safety intended by the certification requirements. The continued structural integrity of aircraft is of concern because factors such as fatigue cracking and corrosion are time‑dependent, and our knowledge about them can best be assessed based on real-time operational experience and the use of the most modern tools of analysis and testing.

In April 1988, a high-cycle transport aeroplane en-route from Hilo to Honolulu, Hawaii, suffered major structural damage to its pressurised fuselage during flight. This accident was attributed in part to the age of the aeroplane involved. The economic benefit of operating certain older technology aeroplanes resulted in the operation of many such aeroplanes beyond their previously expected retirement age. Because of the problems revealed by the accident in Hawaii and the continued operation of older aircraft, both the competent authorities and industry generally agreed that increased attention needed to be focused on the ageing fleet and on maintaining its continued operational safety.

In June 1988, the FAA sponsored a conference on ageing aircraft. As a result of that conference, an ageing aircraft task force was established in August 1988 as a sub-group of the FAA’s Research, Engineering, and Development Advisory Committee, representing the interests of the aircraft operators, aircraft manufacturers, regulatory authorities, and other aviation representatives. The task force, then known as the Airworthiness Assurance Task Force (AATF), set forth five major elements of a programme for keeping the ageing fleet safe. For each aeroplane model in the ageing transport fleet, these elements consisted of the following:

(a)       Select service bulletins describing modifications and inspections necessary to maintain structural integrity;

(b)       Develop inspection and prevention programmes to address corrosion;

(c)       Develop generic structural maintenance programme guidelines for ageing aeroplanes;

(d)       Review and update the supplemental structural inspection documents (SSIDs) which describe inspection programmes to detect fatigue cracking; and

(e)       Assess the damagetolerance of structural repairs.

Subsequent to these five major elements being identified, it was recognised that an additional factor in the Aloha accident was widespread fatigue cracking. Regulatory and industry experts agreed that, as the transport aircraft fleet continues to age, eventually widespread fatigue damage (WFD) is inevitable. Structures Task Groups sponsored by the Task Force were assigned the task of developing these elements into usable programmes. The Task Force was later re‑established as the AAWG of the ARAC. Although there was JAA membership and European operators and industry representatives participated in the AAWG, recommendations for action focussed on FAA operational rules which are not applicable in Europe. It was therefore decided to establish the EAAWG on this subject to implement ageing aircraft activities in Europe, not only for the initial ‘AATF eleven’ aeroplanes, but also other old aircraft and more recently certified ones. EAAWG recommendations followed, leading to the development of guidance material for TCHs and operators, and proposals to develop Sub-part M of JAR OPS. The subsequent establishment of the Agency and new EU regulations led to the current format of Part-M for continuing airworthiness, the associated maintenance programme requirements and to the inclusion of ageing aircraft structures programmes in AMC Part M (M.A.302). AMC 20-20 supported this process and set out means by which TCHs and operators could develop and implement ageing aircraft structures programmes.

This AMC supports DAH and operator compliance with the requirements introduced by Commission Implementing Regulation (EU) 2020/1159 on ageing aeroplane structures, amending Regulation (EU) 2015/640 (Part-26), and the associated CS-26 specifications. The Regulation includes requirements for specific DAHs to perform damage tolerance and other evaluations of existing airframe structure, develop certain data and ICA if they have not already done so, and make it available to operators. Furthermore, operators, in addition to implementing these new ICA as envisaged under Part-M, are required by Part-26 to ensure that approved damage-tolerance-based inspections are obtained and implemented on all repairs and modifications affecting the FCS on aeroplanes certified for 30 passengers or more or for 7 500 lb or more payload.

Points 26.300 through 26.370 of Part-26 provide requirements for a complete retroactively applicable continuing structural integrity programme for specific categories of large aeroplanes. The principal means of compliance with those regulations may be found in CS-26, which, in turn, refers to this AMC.

4.         DEFINITIONS AND ACRONYMS

(a)       For the purposes of this AMC, the following definitions apply:

—      Airworthiness limitation section (ALS) means a section in the instructions for continued airworthiness, as required by points 21.A.61, 21.A.107 and 21.A.120A of Annex I (Part 21) to Regulation (EU) No 748/2012, which contains airworthiness limitations that set out each mandatory replacement time, inspection interval and related inspection procedure.

—      Baseline structure refers to the structure that is designed under the type certificate for that aeroplane model (that is, the ‘as delivered aeroplane model configuration’).

—      Corrosion prevention and control programme (CPCP) is a document reflecting a systematic approach to prevent and to control corrosion in an aeroplane’s primary structure, consisting of basic corrosion tasks, including inspections, areas subject to those tasks, defined corrosion levels and compliance times (implementation thresholds and repeat intervals). A baseline CPCP is established by the type certificate holder, which can be adapted by operators to create a CPCP in their maintenance programme specific to their operations.

—      Damagetolerance (DT) is the attribute of the structure that permits it to retain its required residual strength without detrimental structural deformation for a period of use after the structure has sustained a given level of fatigue, corrosion, and accidental or discrete source damage.

—      Design approval holder (DAH) is the holder of any design approval, including type certificate, supplemental type certificate or earlier equivalent, or repair approval.

—      Damage tolerance data is the combination of DTE documentation and DTI.

—      Damage tolerance evaluation (DTE) is a process that leads to the determination of the maintenance actions necessary to detect or preclude fatigue cracking that could contribute to a catastrophic failure. When applied to repairs and changes, a DTE includes the evaluation of the repair or change and the fatigue-critical structure affected by the repair or change.

—      Damage tolerance inspection (DTI) is a documented inspection requirement or other maintenance action developed by holders of design approvals or third parties as a result of a damage tolerance evaluation. A DTI includes the areas to be inspected, the inspection method, the inspection procedures (including the sequential inspection steps and acceptance and rejection criteria), the inspection threshold and any repetitive intervals associated with those inspections. DTIs may also specify maintenance actions such as replacement, repair or modification.

—      Design service goal (DSG) is the period of time (in flight cycles or flight hours, or both) established at design and/or certification during which the aeroplane structure is expected to be reasonably free from significant cracking.

—      Existing design changes or repairs are changes and repairs which are to be approved before the date of entry into force of this rule.

—      Fatigue-critical alteration structure (FCAS) is equivalent to fatigue-critical modified structure.

—      Fatigue-critical baseline structure (FCBS) is the baseline structure of an aeroplane that is classified by the type certificate holder as a fatigue-critical structure.

—      Fatigue-critical modified structure (FCMS) means any fatigue-critical structure of an aeroplane introduced or affected by a change to its type design and that is not already listed as part of the fatigue-critical baseline structure.

—      Fatigue-critical structure (FCS) is a structure of an aeroplane that is susceptible to fatigue cracking that could lead to a catastrophic failure of the aircraft. For the purposes of this AMC, FCS refers to the same class of structure that would need to be assessed for compliance with JAR 25.571 Change 7 or 14CFR § 25.571(a) at Amendment 25-45, or later. The term ‘FCS’ may refer to fatigue-critical baseline structure, fatigue‑critical modified structure, or both.

—      Inspection start point (ISP) is the point in time when special inspections of the fleet are initiated due to a specific probability of having an MSD/MED condition.

—      Future design changes and repairs are changes and repairs which are to be approved on or after the date of entry into force of this rule.

—      Limit of validity (LOV)(of the engineering data that supports the structural maintenance programme)means, in the context of the engineering data that supports the structural maintenance programme, a period of time, stated as a number of total accumulated flight cycles or flight hours or both, during which it is demonstrated that widespread fatigue damage will not occur in the aeroplane.

—      Multiple-element damage (MED) is a source of widespread fatigue damage characterised by the simultaneous presence of fatigue cracks in similar adjacent structural elements.

—      Multiple-site damage (MSD) is a source of widespread fatigue damage characterised by the simultaneous presence of fatigue cracks in the same structural element.

—      Primary structure is structure that carries flight, ground, crash or pressurisation loads.

—      Published repair data are instructions for accomplishing repairs which are published for general use in structural repair manuals and service bulletins (or equivalent types of documents).

—      Repair assessment guidelines (RAGs) provide a process to establish damage tolerance inspections for repairs on the fuselage pressure boundary structure.

—      Repair assessment programme (RAP) is a programme to incorporate damage‑tolerance-based inspections for repairs to the fuselage pressure boundary structure (fuselage skin, door skin, and bulkhead webs) into the operator’s maintenance and/or inspection programme.

—      Repair evaluation guidelines (REGs)are established by the type certificate holder and guide operators to establish damage tolerance inspections for repairs that affect fatigue-critical structure to ensure the continued structural integrity of all relevant repairs.

—      Structural modification point (SMP) is the point in time when a structural area must be modified to preclude WFD.

—      Widespread fatigue damage (WFD)means the simultaneous presence of cracks at multiple locations in the structure of an aeroplane that are of such size and number that the structure will no longer meet the fail-safe strength or residual strength used for the certification of that structure.

(b)       The following list defines the acronyms that are used throughout this AMC:

5.         CONTINUING STRUCTURAL INTEGRITY PROGRAMME ANDWAY OF WORKING

(a)       General

The programmes and processes described in this and the subsequent paragraphs of this AMC are all part of an acceptable process to provide a continuing structural integrity programme that precludes unsafe levels of cracking.

DAHs and operators are expected to work together to ensure that their continuing structural integrity programmes remain valid.

Points 26.300 through 26.309 provide retroactive requirements for TCHs to establish a continuing structural integrity programme for the type design of large aeroplanes. Aeroplanes certified in compliance with CS-25 Amendment 19 or later have acceptable structural maintenance programmes. Nonetheless, in both cases, there is a need to ensure that the continuing structural integrity programme remains valid throughout the operational life of the aeroplane.

(b)       Maintaining the validity of the continuing structural integrity programme

Point 26.305 requires TCHs to establish a process that ensures that the continuing structural integrity programme remains valid throughout the operational life of the aeroplane, considering service experience and current operations. CS 26.305 (a) and (c) describe the core content of the process required as the means of compliance with point 26.305, and further details are provided in Appendix 5 to this AMC. The intent is for the TCHs of large transport aeroplanes to monitor the continued validity of the assumptions upon which the maintenance programme is based, and to ensure that unsafe levels of fatigue cracking or other structural deterioration will be precluded in service. It should be noted that this requirement applies to all structure whose failure could contribute to a catastrophic failure, and it is not limited to metallic structures or fatigue cracking, but should also encompass composite and hybrid structures.

Typically, large aeroplanes are utilised in well-understood commercial transport scenarios for which conservative or more rational and well-bounded assumptions can be made at the time of certification or when the continuing structural integrity programme is developed. Obvious changes to usage should be addressed for their impact on fatigue and damage tolerance when they occur. In particular, aeroplanes used for conducting surveys, VIP operations, firefighting or other special operations should be considered on a case-by-case basis.

Furthermore, as part of this process, the assumptions made for fatigue, accidental and environmental damage scenarios during certification should, on a regular basis, be validated against service experience to see whether they remain applicable.

The monitoring of operational usage is best achieved in cooperation with the operators, combined with fleet leader sampling inspection programmes. Where data does not correspond to the original certification assumptions, its potential impact on all ageing aeroplane structural programmes and CAW in general must be considered. The degree of impact that a change of usage may have is dependent on the level of conservatism in the selection of the original usage spectrum. It is recommended to review at regular intervals the operational usage data for which a change from the original assumptions would have an impact on the validity of the content of the programme. If this is not done, it might be necessary to investigate the operational usage on each occasion of a service finding in which operational usage could be a contributing factor.

(c)       Way of working

All the ageing aircraft programme elements discussed in this AMC benefit from cooperation between operators and TCHs. The use of structural task groups (STGs) has historically proved very successful in this regard, and is recommended.

On the initiative of the TCH and EASA, an STG may be formed for each aircraft model for which it is decided to put in place an ageing aircraft programme. The STG shall consist of the TCH, selected operator members and EASA representative(s). The objective of the STG is to complete all tasks covered in this AMC in relation to their respective model types, including the following:

—      Develop model specific programmes,

—      Define programme implementation,

—      Conduct recurrent programme reviews as necessary.

It is recognised that it might not always be possible to form or to maintain an STG, due to a potential lack of resources within the operators or TCH. Furthermore, for some mature products, the programmes and their implementation may be sufficiently mature to determine that an STG is not necessary, e.g. when large numbers of aeroplanes have already reached their expected retirement age and none are going to be operated beyond that point. This point could be determined by the LOV, provided that it is not extended. In any case, the responsibilities for ensuring compliance with the applicable requirements are outlined in subparagraph (d) of this paragraph.

An acceptable way of working for STGs is described in the ‘Report on Structures Task Group Guidelines’ that was established by the AAWG with the additional clarifications provided in the following subparagraphs.

(1)       Meeting scheduling

It is the responsibility of the TCH to schedule STG meetings.

(2)       Reporting

The STG may make recommendations for actions via the TCH to EASA. Additionally, the STG should give periodic reports (for information only) to EASA as appropriate with the objective of maintaining a consistent approach.

(3)       Recommendations and decision-making

The decision-making process described in the AAWG Report on Structures Task Group Guidelines paragraph 7 leads to recommendations for mandatory action from the TCH to EASA. In addition, it should be noted that EASA is entitled to mandate safety measures related to ageing aircraft structures, in addition to those recommended by the STG, if it finds it necessary.

(d)       Responsibilities

(1)       The TCH is responsible for developing the ageing aircraft structures programme for each aircraft type, detailing the actions necessary to maintain airworthiness. Other DAHs should develop programmes or actions appropriate to the modification/repair for which they hold approval, unless addressed by the TCH. All the continuing structural integrity programmes, including associated maintenance actions and DTIs, are changes to the ICA and, therefore, are subject to the Part 21 requirements for their promulgation. All DAHs will be responsible for monitoring the effectiveness of their specific programme, and for amending the programme as necessary.

(2)       The operator is responsible for incorporating approved DAH actions necessary to maintain airworthiness into its aircraft-specific maintenance programmes, in accordance with Part-M (point M.A.302) and point 26.370.

(3)       The competent authority of the State of registry, or the continuing airworthiness management organisation (CAMO) when it holds the approval privilege, is responsible for the approval of the aircraft maintenance programme.

(4)       EASA will approve elements of ageing aircraft structures programmes developed by DAHs and may issue ADs to support implementation, where necessary, e.g. to implement applicable inspections and maintenance actions necessary to support the LOV. However, it is intended that Part-M and, where necessary, Part-26 requirements will be the usual means of implementation of ageing aircraft programmes in European registered aircraft.EASA, in conjunction with the DAH, will monitor the overall effectiveness of ageing aircraft structures programmes.

(e)       Continued airworthiness and management of cracks and other damage findings in service

Point 26.305 of Part-26 requires a process to be established that ensures that the continuing structural integrity programme remains valid throughout the operational life of the aeroplane, considering service experience and current operations. One of the elements of this process is the review of new occurrences, existing damage‑tolerance‑based inspections and service bulletins (SBs), which is established in order to determine the need for mandatory changes in cases where inspections alone would not be reliable enough, or to ensure that unsafe levels of cracking are precluded.

For a new type design, the regulations include the damage tolerance approach for preventing catastrophic failures due to fatigue. The damage tolerance approach depends on directed inspection programmes to detect fatigue cracks before they reach their critical sizes.

If an inspection finds cracks in a damage-tolerant fleet, the approval holder, together with EASA, may determine that a demonstrated risk exists, and require additional airworthiness actions, including more rigorous inspection requirements or fleet-wide replacement or modification of the structure.

Cracking is a continued airworthiness issue because cracking usually reduces the strength of the structure to less than its design ultimate strength level. Service history has shown that the reliability of directed inspections is never sufficient to detect all cracks. As the number of crack reports increases, the likelihood that a number of aeroplanes in the fleet have undetected fatigue cracks also increases. Therefore, for areas where fatigue cracks are reported, the likelihood increases that a number of aeroplanes in the fleet will have strengths less than the design ultimate strength level. At some time during operation of the fleet, the likelihood that the strength of any given structure in a fleet is less than the design ultimate strength level may become unacceptably high. The loss of design ultimate strength capability should be a rare event, and EASA rarely knowingly allows the strength of aeroplanes to drop below the design ultimate strength level with any significant frequency.

Approval holders can use the damage tolerance approach to address an unsafe condition. However, it should be understood that damage-tolerance-based inspections may not provide a permanent solution, as explained above, and in cases where cracks are expected to continue to develop in the fleet, the approval holder should propose, and EASA may require, the fleet-wide replacement, modification, or removal from service of the structure.

Other than fatigue crack findings, significant environmental and accidental damage findings should also be taken into account. Initial and critical damage scenarios assumed for certification should be compared to those being reported and where there are differences, the potential airworthiness impact should be evaluated. Differences may include the pattern and extent of cracking, corrosion or accidental damage, the time at which it was discovered and the rate of growth.

More guidance on the continued airworthiness procedures for airframe structures to ensure the validity of the continuing structural integrity programme is provided in Appendix 5.

6.         DAMAGE-TOLERANCE-BASED INSPECTION PROGRAMME

Aeroplanes certified to JAR 25 Change 10 or later or 14 CFR 25 Amdt 54 or later are provided with an airworthiness limitations section (ALS) that includes damage-tolerance-based inspections. Many aeroplanes certified to earlier amendments have also been provided with a DT-based ALS.

Point 26.302 of Part-26 requires TCHs for certain large transport aeroplanes to perform a damage tolerance evaluation (DTE) and establish the associated inspections and other procedures that ensure freedom from catastrophic failures due to fatigue throughout the operational life of the aeroplane. An SSID or ALS developed according to the guidance of this AMC or an SSID mandated under a current EASA AD will satisfy the requirements of point 26.302 of Part-26. In the absence of an approved damage-tolerance-based structural maintenance inspection programme, the TCH, in conjunction with operators, is expected to initiate the development of an SSIP for each aeroplane model. The role of the operator is principally to comment on the practicality of the inspections and any other procedures defined by the TCH and to implement them effectively.

The SSID or ALS should include inspection threshold, repeat interval, inspection methods and procedures. The applicable modification status, associated life limitation and types of operations for which the SSID is valid should also be identified and stated.

For aeroplanes for which a DTE is necessary in accordance with CS 25.571 or point 26.302 of Part-26, all inspections and other procedures must be provided that are anticipated to be necessary throughout the operational life of the aeroplane to prevent catastrophic failures due to fatigue. For an aircraft maintenance programme subject to an LOV under point 26.303 of Part-26 or CS 25.571, the DTE need only provide the inspections and other procedures necessary to prevent catastrophic failures up to the LOV. For other aeroplanes, it is recommended that the ALS includes an LOV or similar limitation on the applicability of the maintenance programme, otherwise the programme should be shown to address the maximum potential usage of the aeroplane based on experience with similar products or a conservative assumption. For an SSIP newly developed to meet point 26.302 of Part-26, the guidance of this AMC applies.

In addition, the inspection access, the type of damage being considered, likely damage sites and details of the resulting fatigue cracking scenario should be included as necessary to support the prescribed inspections.

As a result of a periodic review, the TCH should revise the SSID whenever additional information shows a need. The original SSID will normally be based on predictions or assumptions (from analyses, tests, and/or service experience) of failure modes, time to initial damage, frequency of damage, typically detectable damage, and the damage growth period. Consequently, a change in these factors sufficient to justify a revision would have to be substantiated by test data or additional service information. Any revision to SSID criteria and the basis for these revisions should be submitted to EASA for review and approval of both engineering and maintenance aspects.

7.         DAMAGE TOLERANCE EVALUATION OF REPAIRS AND MODIFICATIONS

Early fatigue or fail-safe requirements (pre-JAR 25.571 Change 7 and 14 CFR §25.571Amdt 45) did not necessarily provide for timely inspection of critical structure so that damaged or failed components could be dependably identified and repaired or replaced before a hazardous condition developed. This applies to modifications and repairs as well as baseline structure.Furthermore, it is known that application of later fatigue and damage tolerance requirements to repairs was not always fully implemented according to the relevant certification bases.

As such, repairs and modifications that have not been subject to a DTE and provided with any necessary DTI may have an adverse effect on the FCS and the safety level achieved by the damage-tolerance-based inspection programme of the baseline structure.

As a result of the above considerations, Part-26 requirements for existing repairs and changes to ageing aeroplane structure were introduced to include specific requirements applicable to certain DAHs and operators of large aeroplanes. Some further details and background are provided here, and Appendix 3 provides additional information on means of compliance with the Part-26 requirements and the associated CS-26 specifications for existing repairs and modifications.

For large aeroplanes with 30 pax or more or having a payload of 3 402 kg (7 500 lb) or more, TCHs must:

(a)       identify and list the FCS according to points 26.306 and 26.307 of Part-26 for FCBS and FCMS respectively and make the list available to assist operators and STCHs needing to identify changes that may require DTE and DTI;

(b)       perform DTE of changes according to point 26.307 of Part-26 and submit the damage tolerance data for approval to EASA; and

(c)       review published repair data and perform DTE in accordance with point 26.308 of Part‑26.

DTIs are ICA and need to be made available to operators according to Part 21. Published repair data includes structural repair manuals (SRMs) and SBs. The data in published repair documentation that needs to be updated includes non-reinforcing repairs such as blending out of scratches, etc. that could be implemented by operators in the future.

For large aeroplanes with 30 pax or more or having a payload of 3 402 kg (7 500 lb) or more, STCHs must:

(a)       identify changes that affect FCBSs and list FCMSs according to point 26.332 of Part-26, and make lists of FCMSs available to assist operators and STCHs needing to identify changes that may require DTE and DTI; and

(b)       perform DTE of changes and published repairs to those changes according to points 26.333 or 26.334 of Part-26 for changes approved on or after 1 September 2003 or before that date respectively, and submit the damage tolerance data to EASA for approval.

CS-26 specifies means of compliance for the DTE itself, and Appendix 3 to this AMC provides means of compliance for the identification of the FCS and implementation of DTI.

The repair evaluation guidelines (REGs) developed by the TCH are intended to assist the operator in addressing the adverse effects of existing reinforcing repairs on the FCS, including the affected adjacent structure, based on damage tolerance principles, consistent with the safety level provided by the SSID or ALS as applied to the baseline structure. In this context, adjacent structure means structure whose fatigue and damage tolerance behaviour and DTE are affected by the reinforcing repair. To achieve this, the REGs should be developed by the TCH and implemented by the operator to ensure that an evaluation is performed of all existing reinforcing repairs to structure that is susceptible to fatigue cracking and could contribute to a catastrophic failure.

Even the best maintained aircraft will accumulate structural repairs when being operated. The AAWG conducted two separate surveys of repairs placed on aircraft to collect data. The evaluation of these surveys revealed that 90 % of all repairs found were on the fuselage, hence these are a priority and repair assessment programmes (RAPs) have already been developed for the fuselage pressure shell of many large transport aeroplanes not originally certified to damagetolerance requirements. 40 % of the repairs were classified as adequate and 60 % of the repairs required consideration for possible additional supplemental inspection during service. Nonetheless, following further studies by the AAWG working groups, it was agreed that repairs to all structure susceptible to fatigue and whose failure could contribute to catastrophic failure should be considered. (Ref. AAWG Report: Recommendations concerning ARAC taskings FR Doc 04-10816 Ref. Aging Airplane safety final rule. 14 CFR 121.370a and 129.16.)

As aircraft operate into high cycles and high times the ageing repaired structure needs the same considerations as the original structure in respect of damagetolerance. Existing repairs may not have been assessed for damagetolerance and appropriate inspections or other actions implemented. Repairs are to be assessed, replaced if necessary or repeat inspections determined and carried out as supplemental inspections or within the baseline zonal inspection programme. A damage-tolerance-based inspection programme for repairs will be required to detect damage which may develop in a repaired area before that damage degrades the load carrying capability of the structure to less than the levels required by the applicable airworthiness standards.

Point 26.309 of Part-26 requires TCHs of aeroplanes with TCs issued prior to 11 January 2008, with 30 pax or more or having a payload of 3 402 kg (7 500 lb) or more, to develop REGs and submit them to EASA for approval.

The REGs should provide data for operators to address existing reinforcing repairs to all structure that is susceptible to fatigue cracking and could contribute to a catastrophic failure. The REGs may refer to the RAP, other existing approved data such as the SRMs and SBs or provide specific means for obtaining data for individual repairs.

This fatigue and damagetolerance evaluation of repairs will establish an appropriate inspection programme or a replacement schedule if the necessary inspection programme is too demanding or not possible. Details of the means by which the REGs and the maintenance programme may be developed are incorporated in Appendix 3 to this AMC.

Point 26.370 of Part-26 directs the operator and organisations responsible for the continuing airworthiness of certain large aeroplanes to revise their maintenance programmes to address the potential adverse effects of repairs and modifications on fatigue-critical structures. The basis for achieving this for existing repairs is the implementation of the REGs supplied by the TCH and, for modifications, the data supplied by the original DAH or a third party contracted by the operator. All repairs and changes that affect the FCS and that are approved and implemented after the applicable date of point 26.370 of Part-26 should be subject to DTE and provided with inspections and other procedures, as necessary. Further guidance on obtaining DTIs and the implementation of the ICA is provided in Appendix 3 to this AMC.

8.         LIMIT OF VALIDITY OF THE MAINTENANCE PROGRAMME AND WIDESPREAD FATIGUE DAMAGE EVALUATION

(a)       Initial WFD evaluation and LOV

All fatigue and damage tolerance evaluations are finite in scope and also therefore in their long-term ability to ensure continued airworthiness. The maintenance requirements that evolve from these evaluations have a finite period of validity defined by the extent of testing, analysis and service experience that make up the evaluation and the degree of associated uncertainties. The limit of validity (LOV) of the engineering data that supports the structural maintenance programme is defined as being not more than the period of time, stated as a number of total accumulated flight cycles or flight hours or both, for which it has been demonstrated that widespread fatigue damage (WFD) is unlikely to occur in the aeroplane structure. To support the establishment of the LOV, the DAH will demonstrate by test evidence, analysis, and, if available, service experience and teardown inspection results of high-time aeroplanes, that WFD is unlikely to occur in that aeroplane up to the LOV. The LOV, in effect, is the operational life of the aeroplane consistent with the evaluations accomplished and the maintenance actions established to prevent WFD.

Note: Although the LOV is established based on WFD considerations, it is intended that all maintenance actions required to address fatigue, corrosion, and accidental damage up to the LOV should be identified in the structural-maintenance programme. All the inspections and other procedures (e.g. modification times, replacement times) that are necessary to prevent a catastrophic failure due to fatigue, up to the LOV, should be included in the ALS of the instructions for continued airworthiness (ICA), as required by CS 25.1529, along with the LOV.

In some cases, the ALS may already contain an LOV which is approved in accordance with a regulation of another authority. There may also be other potentially more restrictive limitations on the validity of the maintenance programme. For these cases, when the TCH needs to publish the LOV as required by point 26.303 of Part-26, this LOV and its relationship with the existing or superseded limitation should be clearly described in order that no operator will exceed the most restrictive applicable limit on the general validity of the maintenance programme.

The likelihood of the occurrence of fatigue damage in an aircraft’s structure increases with aircraft usage. The design process generally establishes a design service goal (DSG) in terms of flight cycles/hours for the airframe. It was generally expected when fatigue and fail-safe rules were first developed that any cracking that occurs on an aircraft operated up to the DSG would occur in isolation (i.e. local cracking), originating from a single source, such as a random manufacturing flaw (e.g. a mis-drilled fastener hole) or a localised design detail. It was considered unlikely that cracks from manufacturing flaws or localised design issues would interact strongly as they grow. The SSIPs described in paragraph 6 ofand Appendix 1to this AMC were intended to find all forms of fatigue damage before they become critical. Nonetheless, it has become apparent that as aircraft have approached and exceeded their DSG, only some SSIPs have correctly addressed WFD as described below.

It should be noted that the majority of aeroplanes in the European fleet are now damage‑tolerance-certified, and that JAR and CS damage tolerance requirements have always required the consideration of all forms of fatigue damage, including damage that would now be described as multiple-site damage (MSD) or multiple-element damage (MED).

JAR 25.571 at Change 7 stated:

‘(b) Damage tolerance (fail-safe) evaluation.

The evaluation must include a determination of the probable locations and modes of damage due to fatigue, corrosion, or accidental damage. The determination must be by analysis supported by test evidence and (if available) service experience. Damage at multiple sites due to prior fatigue exposure must be included where the design is such that this type of damage can be expected to occur.’

AMC 25.571(a), (b) and (e) stated in Section 2.1.1.:

‘d. Provisions to limit the probability of concurrent multiple damage, particularly after long service, which could conceivably contribute to a common fracture path. The achievement of this would be facilitated by ensuring sufficient life to crack initiation.

Examples of such multiple damage are:

i.          A number of small cracks which might coalesce to form a single long crack;

ii.         Failures, or partial failures, in adjacent areas, due to the redistribution of loading following a failure of a single element; and

iii.        Simultaneous failure, or partial failure, of multiple load path discrete elements, working at similar stress levels.

In practice it may not be possible to guard against the effects of multiple damage and failsafe substantiation may be valid only up to a particular life which would preclude multiple damage.’

Nonetheless, it is not clear whether all applicants followed this guidance, hence the development of the Part-26 ageing aeroplane structure requirements and the revision of CS 25.571 at Amendment 19 to include specific means to address WFD. The AMC to these requirements includes the establishment of maintenance actions to modify (or replace) WFD‑susceptible structure prior to the LOV whenever necessary to preclude WFD.

In accordance with point 26.303 of Part-26, TCHs of aeroplanes with MTOMs > 34 019 kg (75 000 lb) have to establish an LOV and the maintenance actions upon which the LOV is dependent, for all model variations and derivatives approved under the TC before 26 February 2021, and all structural changes and replacements to the structural configurations of those aeroplanes that are required by an airworthiness directive (AD) issued before 26 February 2021. Future changes by the TCH to these aeroplanes should also be subject to WFD evaluation. For aeroplane structure certified to CS 25.571 Amdt 19 or later amendment, the fatigue and damage tolerance evaluation requires specific consideration of WFD, see AMC 25.571 paragraph 10.

For a new DTE performed to comply with Part-26 for existing changes or repairs or for new changes or repairs, according to CS 25.571 Amdt 18 or earlier, the evaluation should take into account cracking scenarios that could reasonably be expected to occur in the remaining operational lifetime of the aeroplane in which the repair or modification is implemented. The inspections and other procedures established do not have to include modification and replacement, although the guidance of the applicable ACJ/AMC 25.571 as described above should be considered.

WFD may originate in two basic forms, either as MSD or as MED. With extended usage, uniformly loaded structure may develop cracks in adjacent repetitive features such as fastener holes (MSD), or in adjacent similar structural details (MED). The development of cracks at multiple locations (both MSD and MED) may also result in strong interactions that can affect subsequent crack growth, in which case the predictions for local cracking would no longer apply. An example of this situation may occur at any skin joint where load transfer occurs. Simultaneous cracking at many fasteners along a common rivet line may reduce the residual strength of the joint to the less than the required levels before the cracks are detectable under the maintenance programme established at the time of certification.

Appendix 2 provides guidelines for development of a programme to preclude the occurrence of WFD. Such a programme must be implemented before analysis, tests, and/or service experience indicate that WFD may develop in the fleet. The operator’s role is to provide service experience, to help ensure the practicality of the programme and to ensure it is implemented effectively.

The proposed LOV and the results of the WFD evaluation should be presented for review and approval to EASA for the aeroplane model being considered.

Note: The LOV applies to aeroplanes, not to individual parts. Should there be any concerns about the service life of a removable component containing FCS or principal structural elements (PSEs), a modification or life limitation arising from the WFD evaluation can be mandated on that specific component, which would then need to be tracked.

EASA’s review of the WFD evaluation results will include both engineering and maintenance aspects of the proposal. Per Appendix I to AMC M.A.302, any actions necessary to preclude WFD, including the LOV, are to be incorporated in the maintenance programmes developed in compliance with Part-M. Any SBs or other service information publications revised or issued as a result of in-service MSD/MED findings resulting from implementation of these programmes may require separate AD action.

In the event an acceptable WFD evaluation cannot be completed on a timely basis, EASA may impose service life, operational, or inspection limitations to assure structural integrity of the subject type design.

(b)       Revision of WFD evaluation and LOV

New service experience findings, improvements in the prediction methodology, better load spectrum data, a change in any of the factors upon which the WFD evaluation is based or economic considerations, may dictate a revision to the evaluation. Accordingly, associated new recommendations for service action should be developed including a revised LOV, if appropriate, and submitted to EASA for review and approval of both engineering and maintenance aspects.

An LOV may be extended under the provisions of Part 21. In such cases, the applicant must demonstrate that WFD will not occur in the aeroplane up to the proposed extended LOV. The applicant should consider the age (flight cycles or flight hours or both) of high‑time aeroplanes relative to the existing LOV to determine when to begin developing data to extend it. Because the data is likely to include additional full-scale fatigue testing, the applicant should allow sufficient time to complete such testing and to submit the compliance data for approval. An extended LOV is a major change to the type design of an aeroplane. An extended LOV may also include specified maintenance actions, which would be part of the new LOV approval. Extended LOVs, along with any required maintenance actions for the extended LOV, would be incorporated into the ALS.

Note: Extending an LOV without a physical modification to the aeroplane is not considered a ‘significant’ design change in accordance with point 21.A.101 of Part‑21. However, if extending the LOV requires a physical design change to the aeroplane, the design change is to be evaluated in accordance with point 21.A.101 of Part-21.

For practical purposes, it is suggested that the SRM should also be reviewed and updated to facilitate its continued applicability up to the extended LOV. If this is not done, all SRM‑based repairs will require individual approval. The results together with any necessary actions required to preclude WFD from occurring before the aeroplane reaches the revised LOV should be presented for review and approval by EASA.

Note: Although the extended LOV is established based on WFD considerations, it is intended that all maintenance actions required to address fatigue, corrosion, and accidental damage up to the extended LOV should be identified in the structural‑maintenance programme. All inspections and other procedures (e.g. modification times, replacement times) that are necessary to prevent a catastrophic failure due to fatigue, up to the extended LOV, should be included in the ALS of the ICA, as required by CS 25.1529, along with the extended LOV.

9.         CORROSION PREVENTION AND CONTROL PROGRAMME

A corrosion prevention and control programme (CPCP) is a systematic approach to prevent and to control corrosion in the aircraft’s primary structure. The objective of a CPCP is to limit the deterioration due to corrosion to the level necessary to maintain airworthiness and, where necessary, to restore the corrosion protection schemes for the structure. A CPCP consists of a basic corrosion inspection task, task areas, defined corrosion levels, and compliance times (implementation thresholds and repeat intervals). The CPCP also includes procedures to notify the competent authority and TCH of the findings and data associated with Level 2 and Level 3 corrosion and the actions taken to reduce future findings to Level 1 or better. See Appendix 4 for definitions and further details.

As part of the ICA, the TCH should provide an inspection programme that includes the frequency and extent of the inspections necessary to provide the continued airworthiness of the aircraft. Furthermore, the ICA should include the information needed to apply protective treatments to the structure after inspection. In order for the inspections to be effectively accomplished, the TCH should provide corrosion removal and cleaning procedures and reference allowable limits (e.g. an SRM). The TCH should include all of these corrosion-related activities in a manual referred to as the baseline CPCP. Alternatively, the baseline CPCP may be developed as part of the ICA established by the MRB (ISC) using existing MSG-3 procedures. This baseline CPCP documentation is intended to form a basis for operators to derive a systematic and comprehensive CPCP for inclusion in the operator’s maintenance programme. For operators and owners subject to point 26.370 of Part-26, the operator’s CPCP must take into account the TCH’s baseline CPCP. The competent authority for the operator’s CPCP is the authority responsible for their AMP. The TCH is responsible for monitoring the effectiveness of the baseline CPCP and, if necessary, for recommending changes based on the operator’s reports of findings. In line with the Part-M requirements, when the TCH publishes revisions to their baseline CPCP, these should be reviewed and the operator’s programme adjusted as necessary in order to limit corrosion to Level 1 or better.

The operator should ensure that the CPCP is comprehensive in that it addresses all corrosion likely to affect primary structure, and is systematic in that:

(a)       it provides step-by-step procedures that are applied on a regular basis to each identified task area or zone, and

(b)       these procedures are adjusted when they result in evidence that corrosion is not being limited to an established acceptable level (Level 1 or better).

Note: For an aeroplane with an ALS, in addition to providing a suitable baseline CPCP in the ICA, it is appropriate for the TCH to place an entry in the ALS stating that all corrosion should be limited to Level 1 or better. (This practice is also described in ATA MSG-3.)

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