CS 27.571 Fatigue evaluation of flight structure
ED Decision 2003/15/RM
(a) General. Each portion of the flight structure (the flight structure includes rotors, rotor drive systems between the engines and the rotor hubs, controls, fuselage, landing gear, and their related primary attachments) the failure of which could be catastrophic, must be identified and must be evaluated under sub-paragraph (b), (c), (d), or (e). The following apply to each fatigue evaluation:
(1) The procedure for the evaluation must be approved.
(2) The locations of probable failure must be determined.
(3) In-flight measurement must be included in determining the following:
(i) Loads or stresses in all critical conditions throughout the range of limitations in CS 27.309, except that manoeuvring load factors need not exceed the maximum values expected in operation.
(ii) The effect of altitude upon these loads or stresses.
(4) The loading spectra must be as severe as those expected in operation including, but not limited to, external cargo operations, if applicable, and ground-air-ground cycles. The loading spectra must be based on loads or stresses determined under sub-paragraph (a)(3).
(b) Fatigue tolerance evaluation. It must be shown that the fatigue tolerance of the structure ensures that the probability of catastrophic fatigue failure is extremely remote without establishing replacement times, inspection intervals or other procedures under paragraph A27.4 of appendix A.
(c) Replacement time evaluation. It must be shown that the probability of catastrophic fatigue failure is extremely remote within a replacement time furnished under paragraph A27.4 of appendix A.
(d) Fail-safe evaluation. The following apply to fail-safe evaluation:
(1) It must be shown that all partial failures will become readily detectable under inspection procedures furnished under paragraph A27.4 of appendix A.
(2) The interval between the time when any partial failure becomes readily detectable under sub-paragraph (d)(1), and the time when any such failure is expected to reduce the remaining strength of the structure to limit or maximum attainable loads (whichever is less), must be determined.
(3) It must be shown that the interval determined under sub-paragraph (d)(2) is long enough, in relation to the inspection intervals and related procedures furnished under paragraph A27.4 of appendix A, to provide a probability of detection great enough to ensure that the probability of catastrophic failure is extremely remote.
(e) Combination of replacement time and fail-safe evaluations. A component may be evaluated under a combination of sub-paragraphs (c) and (d). For such component it must be shown that the probability of catastrophic failure is extremely remote with an approved combination of replacement time, inspection intervals, and related procedures furnished under paragraph A27.4 of appendix A.
AMC1 27.571 Fatigue evaluation of flight structure
ED Decision 2023/001/R
This AMC supplements FAA AC 27-1B, § AC 27.571 and should be used in conjunction with that AC when demonstrating compliance with CS 27.571.
(a) Definitions
(1) Rolling contact fatigue (RCF): a form of fatigue that occurs due to the cyclic strains arising from the loading present during rolling contact between two parts of an assembly, e.g. a bearing race and a rolling element.
Note: For the purposes of this AMC, it also includes combinations of rolling and sliding contact phenomena.
(2) Integral race: a bearing race that is an integral part of the transmission structural component such as a gear or shaft.
(b) Explanation
Service experience has shown that RCF can initiate cracks on the surface and below the surface in contact areas structural elements (typically, but not limited to, bearing races and rolling elements and gear teeth) that, in some cases, can propagate to a failure with catastrophic results. It is often assumed that RCF leads first to non-critical partial failures such as micro-pitting and spalling that will be detected before more severe failure modes can develop, such as a complete crack through a part. However, experience has shown that, in some cases, critical failure modes can develop shortly after the occurrence of non-critical partial failures. In such cases, analyses and tests are necessary to demonstrate that sufficient time is available, and the performance of the detection system is adequate to ensure the timely detection to prevent a catastrophic failure.
The certification specifications in CS 27.571 require the identification and fatigue evaluation of the portions of the flight structure, the failure of which could be catastrophic. In order to complete this fatigue evaluation, one of or a combination of the methods proposed in sub-paragraphs (b), (c), (d) and (e) of CS 27.571 should be applied. However, specific characteristics of parts submitted to RCF (e.g. bearings and gears), such as the difficulty to visually inspect the operating nature of these elements, which can lead to mechanical degradation and the impact of RCF, make the application of some of the methods challenging. The procedures of this AMC ensure that the effects of RCF are adequately accounted for in the fatigue evaluations required by CS 27.571.
(c) Procedure
The fatigue evaluation of the portions of the flight structure, the failure of which could be catastrophic, should include, when applicable, the effect of RCF.
For this purpose, steps should be taken to minimise the risk of crack initiation due to RCF on these components (and in particular for integrated bearings races), by minimising contact pressures, specifying high standards for surface finishes, ensuring good lubrication, guaranteeing cleanliness and maintaining lubricant quality regardless of the fatigue evaluation approach selected. The applicant should verify that the selected allowables are suitable to ensure the integrity of the affected components in the operating conditions (temperature, lubrication, cleanliness, etc.) applicable to their design. Experience has demonstrated that it can be beneficial for bearings to be designed so that the reliability of any integrated race subject to the fatigue evaluation is even higher than the least critical race of the bearing. In this way, degradation of the least critical race can lead to detection of the bearing failure before cracking initiates in the integrated race.
As it is difficult to totally preclude cracking initiated by RCF, a ‘fail-safe evaluation’ is recommended wherever possible, such that cracking of affected structural element(s) is detected prior to its residual strength capability falling below the required levels prescribed in CS 27.571(d)(2). Should fatigue cracks initiate and develop into:
(1) Partial failure, such as spalling, the applicant should demonstrate that this condition will be detected at an early stage to avoid catastrophic failure due to further fatigue failure, or loss of integrity of the affected part or any surrounding ones; and
(2) Failure, such as through-cracking of an individual part: the applicant should demonstrate that the remaining structure will withstand service loads and limit or maximum attainable loads (whichever is less) until the failure is detected and damaged components are repaired or replaced to avoid a catastrophic failure.
This demonstration should be performed as appropriate using experience from similar designs, functional tests, structural tests and/or reliable analyses to substantiate that the fail-safe design objective has been achieved, including residual strength demonstration. In addition, the continued safe operation of the affected mechanical system(s) should be ensured for this period considering the potential effect of the failure or partial failure taking into account any pre-existing fatigue damage accrued prior to the failure in the affected component on stiffness, dynamic behaviour, loads and functional performance.
The effectiveness and reliability of means of crack detection for the ‘fail-safe evaluation’, including indirect means of detection such as chip detection systems, and associated instructions for continued airworthiness should be evaluated to show that, if implemented as required, they will result in timely detection and repair or replacement of damaged components. Furthermore, the instructions for continued airworthiness, prescribing the maintenance actions leading up to and following detection of potential failure or partial failure should be substantiated sufficiently to ensure timely repair or replacement of damaged components. The substantiation should consider aspects such as threshold criteria on indicators of means of detection for additional investigative actions and removal from service of the damaged parts, the overall clarity and practicality of the instructions for continued airworthiness and human factors aspects.
In addition to a ‘fail-safe evaluation’, ‘replacement time’ and /or ‘fatigue evaluation’ may be needed in addition to fail-safe evaluation in order to ensure that the assumptions supporting the fail-safety and detection of failure remain valid throughout the operational life of the component.
[Amdt 27/10]
CS 27.573 Damage Tolerance and Fatigue Evaluation of Composite Rotorcraft Structures
ED Decision 2012/021/R
(a) Composite rotorcraft structure must be evaluated under the damage tolerance requirements of sub-paragraph (d) unless the applicant establishes that a damage tolerance evaluation is impractical within the limits of geometry, inspectability, and good design practice. In such a case, the composite rotorcraft structure must undergo a fatigue evaluation in accordance with sub-paragraph (e).
(b) Reserved
(c) Reserved
(d) Damage Tolerance Evaluation:
(1) Damage tolerance evaluations of composite structures must show that Catastrophic Failure due to static and fatigue loads is avoided throughout the operational life or prescribed inspection intervals of the rotorcraft.
(2) The damage tolerance evaluation must include PSEs of the airframe, main and tail rotor drive systems, main and tail rotor blades and hubs, rotor controls, fixed and movable control surfaces, engine and transmission mountings, landing gear, and any other detail design points or parts whose failure or detachment could prevent continued safe flight and landing.
(3) Each damage tolerance evaluation must include:
(i) The identification of the structure being evaluated;
(ii) A determination of the structural loads or stresses for all critical conditions throughout the range of limits in CS 27.309 (including altitude effects), supported by in-flight and ground measurements, except that manoeuvring load factors need not exceed the maximum values expected in service;
(iii) The loading spectra as severe as those expected in service based on loads or stresses determined under sub-paragraph (d)(3)(ii), including external load operations, if applicable, and other operations including high torque events;
(iv) A Threat Assessment for all structure being evaluated that specifies the locations, types, and sizes of damage, considering fatigue, environmental effects, intrinsic and discrete flaws, and impact or other accidental damage (including the discrete source of the accidental damage) that may occur during manufacture or operation;
(v) An assessment of the residual strength and fatigue characteristics of all structure being evaluated that supports the replacement times and inspection intervals established under sub-paragraph (d)(4); and
(vi) allowances for the detrimental effects of material, fabrication techniques, and process variability.
(4) Replacement times, inspections, or other procedures must be established to require the repair or replacement of damaged parts to prevent Catastrophic Failure. These replacement times, inspections, or other procedures must be included in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness required by CS 27.1529.
(i) Replacement times must be determined by tests, or by analysis supported by tests to show that throughout its life the structure is able to withstand the repeated loads of variable magnitude expected in-service. In establishing these replacement times, the following items must be considered:
(A) Damage identified in the Threat Assessment required by sub-paragraph (d)(3)(iv);
(B) Maximum acceptable manufacturing defects and in-service damage (i.e., those that do not lower the residual strength below ultimate design loads and those that can be repaired to restore ultimate strength); and
(C) Ultimate load strength capability after applying repeated loads.
(ii) Inspection intervals must be established to reveal any damage identified in the Threat Assessment required by sub-paragraph (d)(3)(iv) that may occur from fatigue or other in-service causes before such damage has grown to the extent that the component cannot sustain the required residual strength capability. In establishing these inspection intervals, the following items must be considered:
(A) The growth rate, including no-growth, of the damage under the repeated loads expected in-service determined by tests or analysis supported by tests; and
(B) The required residual strength for the assumed damage established after considering the damage type, inspection interval, detectability of damage, and the techniques adopted for damage detection. The minimum required residual strength is limit load.
(5) The effects of damage on stiffness, dynamic behaviour, loads and functional performance must be taken into account when substantiating the maximum assumed damage size and inspection interval.
(e) Fatigue Evaluation:
If an applicant establishes that the damage tolerance evaluation described in sub-paragraph (d) is impractical within the limits of geometry, inspectability, or good design practice, the applicant must do a fatigue evaluation of the particular composite rotorcraft structure and:
(1) Identify structure considered in the fatigue evaluation;
(2) Identify the types of damage considered in the fatigue evaluation;
(3) Establish supplemental procedures to minimise the risk of Catastrophic Failure associated with damage identified in sub-paragraph (e)(2); and
(4) Include these supplemental procedures in the Airworthiness Limitations section of the Instructions for Continued Airworthiness required by CS 27.1529.
[Amdt 27/3]