CS 25.571 Damage tolerance and fatigue evaluation of structure

• ED Decision 2017/015/R found in: CS-25 Amdt 27 - Large Aeroplanes (Jan 2023)

CS 25.571 Damage tolerance and fatigue evaluation of structure ED Decision 2017/015/R (See AMC 25.571) (a) General. An evaluation of the strength, detail design, and fabrication must show that catastrophic failure due to fatigue, manufacturing defects, environmental deterioration, or accidental damage, will be avoided throughout the operational life of the aeroplane. This evaluation must be conducted in accordance with the provisions of subparagraphs (b) of this paragraph, except as specified in subparagraph (a)(4) of this paragraph, for each part of the structure that could contribute to a catastrophic failure. Additionally, a discrete source damage evaluation must be conducted in accordance with subparagraph (e) of this paragraph, and those parts which could contribute to a catastrophic failure must also be evaluated in accordance with subparagraph (d) of this paragraph. In addition, the following apply: (1) The evaluations of subparagraphs (b) and (c) must include: (i) The typical loading spectra, temperatures, and humidities expected in service; (ii) The identification of principal structural elements and detail design points, the failure of which could contribute to a catastrophic failure of the aeroplane; and (iii) An analysis, supported by test evidence, of the principal structural elements and detail design points identified in sub-paragraph (a)(1)(ii) of this paragraph. (2) The service history of aeroplanes of similar structural design, taking due account of differences in operating conditions and procedures, may be used in the evaluations required by this paragraph. (3) Based on the evaluations required by this paragraph, inspections or other procedures must be established as necessary to prevent catastrophic failure, and must be included in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness required by [CS 25.1529](#_DxCrossRefBm1709821031). The limit of validity of the engineering data that supports the structural maintenance programme (hereafter referred to as LOV), stated as a number of total accumulated flight cycles or flight hours or both, established by this paragraph, must also be included in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness. (4) If the results of the evaluation required by subparagraph (b) show that damage tolerance-based inspections are impractical, then an evaluation must be performed in accordance with the provisions of subparagraph (c). If the results of the evaluation show that damage tolerance-based inspections are practical, then inspection thresholds must be established for all principal structural elements and detail design points. For the following types of structure, the threshold must be established based on analyses and/or tests, assuming the structure contains an initial flaw representative of a defect or damage of the maximum probable size that could exist as a result of manufacturing processes or manufacturing or service-induced damage: (i) single load path structure; and (ii) multiple load path ‘fail-safe’ structure and crack arrest ‘fail-safe’ structure, where it cannot be demonstrated that the resulting load path failure or partial failure (including arrested cracks) will be detected and repaired during normal maintenance, inspection, or operation of an aeroplane prior to failure of the remaining structure. (5) Inspection programmes must be established to protect the structure evaluated under subparagraph (b) and (c) against the effects of environmental deterioration and service-induced accidental damage. In addition, a baseline corrosion and prevention control programme (CPCP) must be established. The Airworthiness Limitations Section of the Instructions for Continued Airworthiness must include a statement that requires the operator to include a CPCP in their maintenance programme that will control corrosion to Level 1 or better. (b) Fatigue and damage tolerance evaluation. The evaluation must include a determination of the probable locations and modes of damage due to fatigue, environmental deterioration (e.g. corrosion), or accidental damage. Repeated load and static analyses, supported by test evidence and (if available) service experience, must be incorporated in the evaluation. Damage at multiple sites due to prior fatigue exposure (including special consideration of widespread fatigue damage) must be included in the evaluation where the design is such that this type of damage could occur. An LOV must be established that corresponds to the period of time, stated as a number of total accumulated flight cycles or flight hours or both, for which it has been demonstrated by full-scale fatigue test evidence that widespread fatigue damage will not occur in the aeroplane structure. The type certificate may be issued prior to completion of the full-scale fatigue testing provided that EASA has approved a plan for completing the required tests and analyses, and that at least one calendar year of safe operation has been substantiated at the time of type certification. In addition, the Airworthiness Limitations Section of the Instructions for Continued Airworthiness must specify an interim limitation restricting aircraft operation to not more than half the number of the flight cycles or flight hours accumulated on the fatigue test article, until such testing is completed, freedom from widespread fatigue damage has been established and the LOV is approved. The extent of damage for residual strength evaluation at any time within the operational life of the aeroplane must be consistent with the initial detectability and subsequent growth under repeated loads. The residual strength evaluation must show that the remaining structure is able to withstand loads (considered as static ultimate loads) corresponding to the following conditions: (1) The limit symmetrical manoeuvring conditions specified in CS 25.331 at all speeds up to VC and in CS 25.345. (2) The limit gust conditions specified in [CS 25.341](#_DxCrossRefBm1709820927) at the specified speeds up to VC and in CS 25.345. (3) The limit rolling conditions specified in [CS 25.349](#_DxCrossRefBm1709820948) and the limit unsymmetrical conditions specified in [CS 25.367](#_DxCrossRefBm1709820978) and [CS 25.427(a) through (c)](#_DxCrossRefBm1709820945), at speeds up to VC. (4) The limit yaw manoeuvring conditions specified in [CS 25.351](#_DxCrossRefBm1709820961) at the specified speeds up to VC. (5) For pressurised cabins, the following conditions: (i) The normal operating differential pressure combined with the expected external aerodynamic pressures applied simultaneously with the flight loading conditions specified in sub-paragraphs (b)(1) to (b)(4) of this paragraph if they have a significant effect. (ii) The maximum value of normal operating differential pressure (including the expected external aerodynamic pressures during 1 g level flight) multiplied by a factor of 1·15 omitting other loads. (6) For landing gear and other affected airframe structure, the limit ground loading conditions specified in [CS 25.473](#_DxCrossRefBm1709820981), [CS 25.491](#_DxCrossRefBm1709821016) and [CS 25.493](#_DxCrossRefBm1709821020). If significant changes in structural stiffness or geometry, or both, follow from a structural failure, or partial failure, the effect on damage tolerance must be further evaluated. (c) *Fatigue (safe-life) evaluation*. Compliance with the damage-tolerance requirements of sub-paragraph (b) of this paragraph is not required if the applicant establishes that their application for particular structure is impractical. This structure must be shown by analysis, supported by test evidence, to be able to withstand the repeated loads of variable magnitude expected during its service life without detectable cracks. Appropriate safe-life scatter factors must be applied. Until such time as all testing that is required for compliance with this subparagraph is completed, the replacement times provided in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness may not exceed the total accumulated flight cycles on the test article test life divided by the applicable scatter factor. (d) *Sonic fatigue strength.* It must be shown by analysis, supported by test evidence, or by the service history of aeroplanes of similar structural design and sonic excitation environment, that: (1) Sonic fatigue cracks are not probable in any part of the flight structure subject to sonic excitation; or (2) Catastrophic failure caused by sonic fatigue cracks is not probable assuming that the loads prescribed in sub-paragraph (b) of this paragraph are applied to all areas affected by those cracks. (e) *Discrete source* damage tolerance *evaluation.* The aeroplane must be capable of successfully completing a flight during which likely structural damage occurs as a result of bird impact as specified in [CS 25.631](#_DxCrossRefBm1709821047). The damaged structure must be able to withstand the static loads (considered as ultimate loads) which are reasonably expected to occur at the time of the occurrence and during the completion of the flight. Dynamic effects on these static loads do not need to be considered. Corrective action to be taken by the pilot following the incident, such as limiting manoeuvres, avoiding turbulence, and reducing speed, may be considered. If significant changes in structural stiffness or geometry, or both, follow from a structural failure or partial failure, the effect on damage tolerance must be further investigated. [Amdt 25/18] [Amdt 25/19]