described in AC 27-1B MG 11, has proven to be successful in providing an adequate level of reliability for transport category rotorcraft. However, it was recognized in the 1980’s that higher levels of reliability might be realized by taking into account the fatigue strength-reducing effects of damage that experience has shown can occur in manufacture or in operational service. The introduction of composites led the manufacturers and regulatory authorities to develop a robust Safe-Life methodology by taking into account the specific static and fatigue strength-reducing effects of aging, temperature, moisture absorption, impact damage, and recognition of an accepted industry standard. Furthermore, where clearly visible damages resulted from impact or other sources, inspection programs were developed to maintain safety. In parallel, crack growth methodology has been successfully used for solving short-term airworthiness problems in metallic structures of rotorcraft, and as the certification basis for civil and military transport aircraft applications. These advances in design, analytical methods, and industry practices made it feasible to address certain types of damage, which could result in fatigue failure. Consistent with this, the regulatory requirements of § 29.571 were substantially revised by Amendment 28. While many years have passed since its introduction, Amendment 28 has had little exposure to use for certification of completely new rotorcraft designs. However, the general understanding of rotorcraft fatigue tolerance evaluation has developed considerably in the interim and an additional amendment was determined to be appropriate. The latest Amendment 29- 55 of part 29 and the associated revisions to advisory material were introduced to improve the currency and understanding of the rule and clarify the differing approaches and methods available for accomplishing fatigue tolerance evaluation of rotorcraft metallic structure.
(2) This guidance provides material with respect to the fatigue tolerance requirements for metallic structure and is supplemented by AC 27-1B MG 11 for evaluations using the Safe-Life methodology and other general fatigue considerations.
d. Introduction.
(1) Definitions. The following definitions are applicable when used within the context of this guidance material.
(i) As-manufactured structure is a structure that passes the applicable quality control process and has been found to conform to an approved design within the allowable tolerances.
(ii) Barely Detectable Flaw (BDF) is the worst-case flaw that is expected to remain on the structure for its operational life.
(iii) Catastrophic failure is an event that could prevent continued safe flight and landing.
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(iv) Clearly Detectable Flaw (CDF) is the worst-case detectable flaw that would not be expected to remain in place for a significant period of time without corrective action.
(v) Damage is a detrimental change to the condition of the structure or assembly. In the context of this guidance material it is used as a generic term to describe all types of flaws including those caused by environmental effects and accidental damage arising in manufacture, maintenance or operation.
(vi) Damage Tolerance is the attribute of the structure that permits it to retain its required residual strength without detrimental structural deformation for a period of un-repaired use after the structure has sustained a given level of fatigue, corrosion, accidental, or discrete source damage.
(vii) Discrete flaw is a flaw that is not inherent in the design and is caused by an external action, such as corrosion, scratches, gouges, nicks, fretting, wear, impact, and potentially cracks initiated by fatigue.
(viii) Fatigue is a degradation process of a structure subject to repeated loads that may involve four phases (e.g., nucleation of many micro-cracks, coalescence of some micro-cracks to one major macro-crack, stable crack growth, unstable crack growth) and immediate failure. The boundaries between these phases are, in practice, not always easily defined. Crack initiation methods (e.g., using the S -N curve and the Miner’s Rule) are generally used to address the first two phases. Linear Fracture Mechanics methods (e.g., using da/dn - ΔK and fracture toughness data) are generally used for the latter two phases.
(ix) Fatigue Loads are repeated loads, which induce a repeated variation of stress versus time in a structure.
(x) Fatigue Tolerance is the ability of a structure, either in an as-manufactured or damaged condition, to tolerate specified operational loading for a given period of use without initiating cracks, and assuming they initiate, tolerate their growth, without failure, under specified residual strength loads.
(xi) Flaw is an imperfection, defect, or blemish and may be either discrete or intrinsic.
(xii) Inspection interval is the maximum period of usage allowed for a structure between inspections. At the end of this period, the structure is inspected and if there is no damage detected, the structure may be returned to service for another inspection interval.
(xiii) Intrinsic flaw is a flaw that is inherent in the design and manufacture of the part, situated within it or peculiar to it, such as inclusions, cracks, forging laps, or porosity.
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(xiv) Limit Loads are the maximum loads to be expected in service, as defined in § 29.301(a).
(xv) Multiple Load Path is identified with a redundant structure of multiple and distinct elements, in which the applied loads would be safely redistributed to other load carrying members after complete failure of one of the elements. These may be Active, where two or more elements are loaded during operation to a similar load spectrum, or Passive, where one or more of elements of the structure are relatively unloaded until failure of the other element(s).
(xvi) Principal Structural Elements (PSE) are structural elements that contribute significantly to the carrying of flight or ground loads and the fatigue failure of which could result in catastrophic failure of the rotorcraft.
(xvii) Residual Strength is the level of strength retained by a structure with damage present.
(xviii) Retirement (Replacement) Time of a component is that number of events such as flight hours or landings at which the part must be removed from service regardless of its condition.
(xix) Safe-Life is the number of events, such as flight hours or landings, for a structural component during which there is a low probability that the strength will degrade below its design ultimate value due to fatigue damage initiating cracks.
(2) General. The objective of fatigue tolerance evaluation is to prevent catastrophic failure of the structure by mitigation of the effects of damage in combination with fatigue throughout the life of the rotorcraft.
(i) Fatigue tolerant design as substantiated by fatigue tolerance evaluation methods such as those outlined in this guidance is required for all PSE’s, unless it entails such complications that an effective structure that is tolerant to damage cannot be achieved within the limitations of geometry, inspectability, or good design practice.
In such cases, the particular type of damage at issue must be identified and alternative measures should be taken to minimize both the risk of acquiring that damage and its consequences.
(ii) To perform an evaluation first requires an understanding of the potential threats (resulting in damage) that may modify the fatigue behavior of the component.
The principal concerns of this guidance are consideration of all damage sources and of the fatigue loads and rotorcraft usage. Further mitigation of the sources of damage may be achieved by adoption of a critical parts plan to help ensure that the condition of the part remains as envisaged by the designer throughout its life cycle (see § 29.602).
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(iii) The need for the use of complex inspection techniques or equipment or highly trained personnel (resources that may not be available to the small operator or in remote areas of operation) should be considered when establishing the methodology.
When inspections cannot be relied upon for detection of small cracks or other damage, then retirement times must be established that account for the probable types and locations of the damage, including consideration of cracks.
(iv) A retirement time should be provided for all components, including those subject to inspection, whose fatigue behavior is not reliably established to a point well beyond the life of the rotorcraft. This is intended to prevent the continued use of components beyond the point that ultimate load capability may no longer be assumed to exist in the rotorcraft due to the onset of fatigue cracking. This is particularly important for single load path components or a structure prone to widespread fatigue damage.
(v) Experience with the application of methods of fatigue tolerance evaluation indicates that a relevant test background should exist in order to achieve the design objective. It is general practice within industry to conduct tests to obtain design information and for certification purposes. Damage location, fatigue characteristics, and crack growth data based on test results and service history of similar parts, if available, should be considered when establishing inspections and retirement times.
The FAA should agree upon the extent of supporting evidence necessary for each phase of the evaluation process outlined below.
(3) Essential Considerations. In order to satisfy the requirements of § 29.571, consideration should be given to the following issues in order to demonstrate compliance.
(i) Selection of PSE. All structure, structural elements, and assemblies, the failure or undetected failure of which could result in catastrophic failure of the rotorcraft, should be identified as PSE [see paragraph f.(2)]. To do this, a failure mode and effects analysis or similar method may be used. Specific areas of interest within the PSE that may require particular attention include the following:
(A) irregularly shaped parts, or those containing numerous or superimposed fillets, holes, threads, or lugs;
(B) parts of unique design for which no past service experience is available;
(C) new materials or processes for which there is no previous experience;
(D) bolted or pinned connections;
(E) parts subject to fretting;
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(F) complex casting; and (G) welded sections.
(ii) In-flight measurement to determine the loads or stresses (steady and oscillatory) for the PSEs in all critical conditions throughout the range of limitations in § 29.309 (including altitude effects), except that maneuvering load factors need not exceed the maximum values expected in operations. See paragraph f.(3).
(iii) Loading spectra as severe as those expected in operation including external load operations, if applicable, and other high frequency power cycle operations. See paragraphs f.(3) and f.(4).
(iv) A threat assessment of probable damage, including a determination of the probable locations, types, and sizes should be performed. In particular, the assessment should include an evaluation of the details of the specific work processes used on each component, operational environment, and maintenance practices to determine the potential for damage. See paragraph f.(5).
(v) Inspectability of the rotorcraft, inspection methods, and detectable flaw sizes should be compatible with the chosen fatigue tolerance methods and validated by trials conducted under realistic conditions. See paragraph f.(6).
(vi) For each PSE, one or more fatigue tolerance methodologies should be selected to ensure each specific damage resulting from the threat assessment is addressed and to satisfy the requirement for inspections and retirement times as discussed in paragraph e. of this guidance. The fatigue tolerance characteristics (including variability) of the structure and materials therein should be evaluated as necessary to support the evaluation. Generally, this will include understanding the fatigue strength, fatigue crack propagation characteristics of the materials used, and the structure and the residual strength of the damaged structure. See paragraphs e., f.(7) and f.(8).
(vii) Fatigue Tolerance Results of the evaluation should be used to provide data in the Limitations Section of the Instructions for Continued Airworthiness. See paragraph f.(9).
e. Fatigue Tolerance Evaluation. A fati gue tolerance evaluation, by analysis and tests, of the PSE is required to establish inspections and retirement times, or approved equivalent means, to avoid catastrophic failure due to fatigue cracking during the operational life of the rotorcraft. The evaluation should consider the impact of the probable threats identified on the fatigue performance and residual strength of all critical areas of each PSE. A number of different fatigue evaluation methods have evolved over the years. Seven of these methods are recognized and discussed in detail in this guidance. The seven methods are summarized as a table in Figure AC 29.571B -1.
Also noted in the table is the safety management strategy the specific method supports, Page C - 103
AC 29-2C, Chg 4 12/2/2011 the analysis category in which they belong, and whether the specific method can be used to address the types of damage identified in the threat assessment.
(1) Each approach results in information that can be used to support establishment of retirement times or inspection requirements. Four methods are used to support safety-by-retirement strategies and they result in retirement times. The other three methods are used to support safety-by-inspection strategies and the result is in-service inspection requirements.
(2) In some cases, application of one method may be sufficient to achieve acceptable fatigue tolerance. In other cases more than one method may be needed.
For example, use of Safe-Life Retirement in combination with Crack Growth Inspections could be an effective way to manage fatigue due to all possible sources.
(3) All the methods listed, with the exception of Safe-Life Retirement, were developed to explicitly address some level of damage. All the methods can theoretically be implemented analytically or by test. However, some of the methods are more practically implemented analytically and some are best implemented by test.
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Method PARAGRAPH Strategy Analysis Category Threat Assessment Results Safe-Life Retirement e.(6)(i)(A) Retire Crack Initiation Not Included Safe-Life Retirement with BDF(s)
e.(6)(i)(B) Retire Crack Initiation Not Including Cracks Safe-Life Retirement with CDF(s)
e.(6)(i)(C) Retire Crack Initiation Not Including Cracks Safe-Life Inspection for CDF(s)
e.(6)(i)(D) Inspect Crack Initiation Included Safe-Life Inspection for a failed element
e.(6)(i)(E) Inspect Crack Initiation Included if Considered for all Elements Crack Growth Retirement
e.(6)(ii)(A) Retire Crack Growth Included if Crack Bounds Damage Crack Growth Inspection
e.(6)(ii)(B) Inspect Crack Growth Included Figure AC 29.571B-1: Seven Fatigue Evaluation Methods discussed in this guidance
(4) From an analytical standpoint, these methods fall into one of two categories, crack initiation or crack growth. Each of the seven methods is briefly described below in paragraphs e.(6)(i) and e.(6)(ii), depending on the category.
(5) In-service experience may be used to support establishing fatigue tolerance characteristics when it is shown on a similar structure.
(6) Fatigue Evaluation Methods.
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(i) Crack Initiation Methods. The methods described in this section are categorized as crack initiation methods since they involve quantifying the time it takes for a crack to initiate at a critical area in an as-manufactured part or at a critical area that has sustained some level of damage. Analytically these methods depend on fatigue data (e.g., stress versus number of cycles (S-N) curves) and cumulative fatigue damage algorithms (e.g., Miner’s Rule) to establish a high margin retirement time.
Testing that supports these methods employs specimens that are as -manufactured or ones that have been preconditioned with damage as identified in the threat assessment.
(A) Safe-Life Retirement. Safe- Life Retirement is a crack initiation method that accounts for damage induced by fatigue loading but does not account for flaws and defects due to manufacturing and in-service conditions. Application of this method results in a replacement time based on the time to initiate a crack in an as - manufactured part. Analysis or tests may be used to determine the crack initiation life.
The rationale behind this method is based on part replacement before the probability of initiating a crack becomes significant. This method needs to be supplemented by other methods to account for damage. For compliance details, see paragraph f.(7)(i).
(B) Safe-Life Retirement with a Barely Detectable Flaw (BDF). Safe- Life Retirement with a BDF is a crack initiation methodology that explicitly addresses the effect of damage that is considered barely detectable and is therefore likely to go unnoticed for the life of the part. Application of this method results in a replacement time based on the time to initiate a crack from a BDF. Analysis or tests may be used to determine the crack initiation life. The rationale behind this method is based on part replacement before the probability of initiating a crack is significant. Damage in excess of the BDF must be addressed using other methods. For compliance details, see paragraph f.(7)(ii).
(C) Safe-Life Retirement with a Clearly Detectable Flaw (CDF). Safe- Life Retirement with a CDF is a crack initiation methodology that explicitly addresses the effect of damage that is considered clearly detectable but conservatively recognizes that it would remain in place without corrective action prior to the retirement time of the part. Application of this method results in a retirement time based on the time to initiate a crack from a CDF. Analysis or tests may be used to determine the crack initiation life.
The rationale behind this method is based on part replacement before the probability of initiating a crack is significant. Use of this method by itself could achieve acceptable fatigue tolerance and may preclude the need for any mandated directed inspections.
See paragraph f.(7)(iii) for compliance details.
(D) Safe-Life Inspection for a CDF. Safe- Life Inspection for a CDF is a crack initiation method that explicitly addresses the effect of damage that is considered clearly detectable and would therefore not be expected to remain in place without corrective action for any significant period of time. Application of this method results in a directed inspection task with an interval based on the time to initiate a crack from a clearly detectable flaw. Analysis or tests may be used to determine the crack initiation Page C - 106
life. The rationale behind this method is based on visual detection and disposition of the flaw before the probability of initiating a crack is significant. Damage that is not detectable must be addressed by other methods and the cumulative effects of fatigue prior to and following the advent of the damage should be considered. For compliance details, see paragraph f.(8)(i).
(E) Safe-Life Inspection for a failed element. Safe- Life Inspection for a failed element is a crack initiation method. It results in an inspection for a completely failed load path with an interval based on the crack initiation life of the adjacent structure accounting for internal load redistribution due to failure of the load path that is to be inspected. This method can only be applied if the structure is initially designed for limit load capability with the failed element. The rationale behind this method is based on visual detection and disposition of the failed load path before the probability of initiating a crack in the adjacent structure becomes significant. Therefore it may not be appropriate if the damage that has led to the failure of the first load path could similarly affect the remaining path. For compliance details, see paragraph f.(8)(iii).
(ii) Crack Growth Methods. The methods described in this paragraph are categorized as crack growth methods since they involve quantifying the time it takes a crack at a critical area to grow from some initial size to some final size. Analytically these methods depend on crack growth rate properties (e.g., da/dN vs. ΔK vs. R) and fracture properties (e.g., KIC). Using these properties, Fracture Mechanics based tools are used to predict crack growth and final fracture. Testing that supports these methods employs specimens that contain cracks and involves close monitoring to document actual crack growth and final fracture.
(A) Crack Growth Retirement is a crack growth method that explicitly addresses the largest damage that could occur during manufacture or operation of the rotorcraft. This damage is modeled as a bounding equivalent crack (BEC) established based on the results of the threat assessment. Application of this method results in a retirement time based on the time for the initial crack to grow large enough to reduce the residual strength to design limit level. Since typical BECs are relatively small and thus difficult to induce in test specimens, this method is typically implemented analytically. The rationale behind this method is based on part retirement before the largest probable damage, modeled as a crack, would reduce the residual strength below design limit. Use of this method by itself could achieve acceptable fatigue tolerance and preclude the need for any mandated inspections provided all threats are accounted for by the BECs. For compliance details, see paragraph f.(7)(iv) (B) Crack Growth Inspection is a crack growth method that explicitly addresses damage that could occur during manufacture or operation of the rotorcraft.
An in-service inspection method is selected that defines a detectable crack size, which could be as large as a completely failed load path. An inspection interval is established based on the time for the detectable crack to grow to critical size or for the residual strength of the adjacent structure to drop to design limit due to continuing crack growth in it. This method is applicable to single or multiple load path structure and inspection Page C - 107
AC 29-2C, Chg 4 12/2/2011 for a completely failed load path or less. This method may be addressed by analysis supported by test depending on the difficulty of introducing into the specimen the inspectable crack or failed load path. The rationale behind this approach is based on detection and disposition of a crack or failed load path before residual strength is reduced below the design limit load. For compliance details, see paragraph f.(8)(ii).
f. Means of Compliance.
(1) General. The results of the fatigue tolerance evaluation required by § 29.571 are used to establish operational procedures that are meant to minimize the risk of catastrophic failures during the operational life of the rotorcraft. It is required that the evaluation performed considers the effect of damage that could result from potential threats present during manufacture and operation. An assessment of probable threats is required to identify the damage that must be considered in the fatigue tolerance evaluation.
(i) The fatigue tolerance evaluation should establish both retirement times and inspection intervals, or approved equivalent means, to prevent any catastrophic failures. Retirement times should be set to ensure that baseline ultimate strength capability is not compromised for as-manufactured structures and structures where the damage is likely to be undetected during the operational life. Intervals for inspections for detectable damage must be established so that strength capability will never fall below maximum design limit level. The intent is that if damage does occur, the structure will retain the capability to withstand reasonable loads without catastrophic failure or excessive structural deformation until the damage is detected and the structure is replaced or repaired. If inspections cannot be established within the limitations of geometry, inspectability, or good design practice, then supplemental procedures, when available, should be established that would minimize the risk of damage being present or leading to a catastrophic failure.
(ii) The following considerations will assist the successful design of a fatigue tolerant structure.
(A) Use multiple-element and multiple load path construction with provisions for crack stoppers that can limit (arrest) the growth of cracks while maintaining adequate residual strength.
(B) Select materials and stress levels that preclude crack growth or crack initiation from flaws or that provide a controlled slow rate of crack propagation combined with high residual strength after initiation of cracks. Test data should substantiate material properties.
(C) Design for detection of damage (i.e., cracks and flaws) and retirement or repair.
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(D) Provide provisions that limit the occurrence of damage and the probability of concurrent multiple damage, particularly after long service.
(iii) Section 29.571 requires that the applicant's proposed compliance methodology must be submitted and be approved by to the Administrator. Therefore, the applicant should coordinate the involvement of the FAA from an early stage. The proposed means of compliance should include the following items.
(A) A list of PSEs to be evaluated.
(B) The results of threat analyses for each PSE including type, location, and size of the damage that will be considered in order to establish retirement times, inspections, or other procedures.
(C) Inspection criteria that includes an estimate of detectability or inspectability, along with any supplemental procedure to minimize the risk of damage.
(D) The analysis methods and supporting test data that will establish retirement times, inspections, or other procedures.
(2) Identification of PSE. The fatigue tolerance evaluation should first consider all airframe structure and structural elements, and assemblies in order to identify the PSE. The structural elements and assemblies identified as PSE should be formally submitted to the FAA with justification for the PSE based on good design practice, service history with similar structure, drawing reviews, static analysis issues, or other appropriate means.
(i) A Failure Mode and Effects Analysis or similar method may be used to identify structures whose failure due to fatigue can lead to catastrophic failure of the rotorcraft. The need to design a PSE for fatigue tolerance when they are supplied by third parties (e.g., actuators) should be clearly identified in the rotorcraft manufacturer’s specification for the part. The list of PSE will likely include structural elements and assemblies that will be subjected to significant fatigue loading expected during the operational life of the rotorcraft. This may include the following rotorcraft parts:
(A) Rotors: blades, hubs, hinges, attachment fittings, vibration dampening devices;
(B) Rotor drive systems (parts connecting rotors to engines): gears, shafts, gear housings, couplings;
(C)Rotor control systems: actuators, pitch control system, swashplate, servo flaps;
(D) Fuselage (airframe): rotor system support structure, landing gear attachment;
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AC 29-2C, Chg 4 12/2/2011 (E) Fixed and movable control surfaces: stabilizer;
(F) Engine, transmission or equipment mountings: APU, auxiliary gearbox;
(G) Landing gear;
(H) Folding systems: main blade, tail beam.
(ii) Analyses and fatigue tests on complete structures or representative subelement structures can determine the locations within PSE that need to be identified for fatigue tolerance evaluation. The following should be considered:
(A) Strain gauge data on undamaged structure that can identify high stress points.
(B) Analysis that shows high stress or small margin of safety values.
(C) Locations where permanent deformation occurred in static tests.
(D) Locations where failure has occurred in as-manufactured structure fatigue tests.
(E) Locations where the potential for fatigue damage has been identified by analysis.
(F) Locations where the maximum allowed stress occurs when an adjacent element fails.
(G) Locations in structure needed to maintain adequate residual strength that has high stress concentration values.
(H) Locations where detection would be difficult.
(I) Locations where service experience with similar components indicates potential for fatigue or other damage (e.g., fretting, corrosion, wear).
(3) Flight Loads Measurement Program. The simulation of expected spectrum loads for each PSE should be based on flight recorded strain gauge data collected as part of a structured flight test program. The PSE spectrum loads include the steady state, transient, and vibratory loads that are expected in operation. AC 27- 1B MG 11, provides further detail for development and use of flight measured loads as the basis for spectrum loads used in the fatigue tolerant evaluations.
(4) Rotorcraft Usage Spectrum.
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