AC 29-2C, Chg 4 12/2/2011
(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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(i) The usage and loading spectrum should be developed so that it is unlikely that the actual usage and loads will cause fatigue damage or crack growth rates beyond those associated with the defined spectrum used in the fatigue tolerance evaluation.
The usage spectrum allocating percentage of time or frequencies of occurrence to flight conditions or maneuvers should be based on the expected usage of the rotorcraft.
Considerations should include flight history, recorded flight data, design limitations established in static strength requirements, and recommended operating conditions and limitations specified in the rotorcraft flight manual.
(ii) The fatigue load spectrum developed for fatigue testing and analysis purposes should be representative of the anticipated service usage. Low amplitude load levels that can be shown not to contribute fatigue damage may be omitted (truncated). Simplification of the spectrum loads may also include summing (binding) of percent times or cycles with common steady and vibratory load values.
(iii) The steady state, transient, and vibratory flight load assigned to each regime in the spectrum and utilized in the fatigue tolerance evaluations for each condition should take into account combinations of altitude, center of gravity (CG), gross weight (GW), airspeed, etc., considered to be representative of expected G W/CG mission configurations.
(iv) The usage spectrum should be presented to the FAA for their concurrence. It should include normal operation over the range of rotorcraft configurations including a percent time under ‘external load’ conditions. This spectrum should represent a “composite worst-case” compilation that includes all of the critical conditions that the rotorcraft is expected to experience during performance of the design missions.
(v) AC 27-1B MG 11, provides further detail for the development of the usage spectrums used in the fatigue tolerance evaluations.
(5) Threat Assessment.
(i) A determination should be made of all potential threats that could occur during the manufacturing and service life that may cause damage to each PS E. A threat assessment should be performed for each PSE. To acquire sufficient knowledge of the component and of its global environment, the following items must be identified:
(A) manufacturing process.
(B) quality control process.
(C) prescribed storage, transport, handling, assembly and maintenance aspects of the component, and of the surrounding components .
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AC 29-2C, Chg 4 12/2/2011 (D) operational environment.
(E) potential for corrosion including that from contamination by corrosive fluids.
(F) potential for impact damages from debris, dropped tools, hail, tramping underfoot during maintenance, etc.
(G) potential for wear.
(ii) To determine types, locations, and sizes of the probable damages, considering the time and circumstances of their occurrence, the following should be considered:
(A) Intrinsic flaws and other damage that could exist in an as-manufactured structure based on the evaluation of the details and potential sensitivities involved in the specific manufacturing work processes used.
(B) Damage that could be expected to occur during prescribed activities associated with storage, transport, handling, assembly, maintenance, overhaul, repair and operation of the component and of the surrounding components including impac ts, scratches, fretting, corrosion, contamination, wear, and loss of bolt torque.
(C) Previous experience and data collected on similar events and on similar components; materials, and processes should be considered in identifying risks and causes of damages and their effects in inducing flaws or cracks.
(D) Metallurgical evaluations, manufacturing records and overhaul and repair reports, field service reports, incident and accident investigations, and engineering judgment may be used as supporting data.
(E) When data are not available, the threat should be experimentally simulated and the effect established through tests and analysis. With agreement of the FAA, an upper cut-off value may be established for each class of damage.
(F) Credit may be given to manufacturing, transport, handling, installation, and maintenance instructions finalized to minimize or avoid damages. Examples of these processes or instructions could be: "frozen manufacturing processes," Flight Critical Parts programs, material selection to mitigate intrinsic flaws like inclusions and defects, procedures to reduce deviations from nominal structures, etc.
(G) Credit may be given to protection of structures, such as the use of protective coatings, shielding and plating against corrosion, fretting, and impacts.
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(H) Critical areas will be assumed as a typical location of the damage, unless proper justification is provided to limit the applicability to specific areas or sections of the part.
(iii) Classification of Damage.
(A) The results of the threat assessment are used to classify the damage used in the fatigue tolerance evaluation. The process employed to classify the damage will depend on the fatigue tolerance evaluation method to be used. Depending on the method, a BDF, a CDF, a BEC, or an initial inspectable crack must be established.
(B) For each damage type identified, the sizes to be considered should be representative of the maximum sizes that might not be detected by the inspection techniques established for the component. Sizes exceeding those that are likely to occur do not need to be considered. Standard sizes of damage or standard level of aggression may be derived from previous experience. Each applicant will be required to present justification for damage and crack sizes to be used in the fatigue tolerance evaluations. Within the operational life, defect sizes that have been found in service should be correlated with the sizes used in the design certification.
(C) Barely Detectable Flaw (BDF). For retirement time analysis, flaw sizes that are “barely detectable” may be used to conservatively represent the worst case of undetectable flaws. Alternatively, when the detectable size is larger than the one identified by the threat assessment, a smaller size, but one not less than the flaw size likely to occur, can be used. Sometimes an “allowable” detectable size is established as acceptable for a specific manufacturing process, such as castings, to remain in place for the life of the structure. When it is impossible to simulate that maximum allowable size in the test specimen, the sizes available in the specimen may be used, provided the subsequent analysis of the test result conservatively accounts for the shortfall in the damage size.
(D) Clearly Detectable Flaw (CDF). For inspection intervals, flaw sizes that are “clearly detectable” may be used. The largest discrete size of a CDF to be considered may be limited to the maximum size of the CDF that is likely to remain in place for a significant period of time and not be detected during routine inspections for general conditions and normal observations by knowledgeable personnel. The damage size used may be limited to the maximum probable size identified in the threat assessment. For multiple load path structure, the number of failed load paths to be considered should be established.
(E) Bounding Equivalent Crack (BEC). A Bounding Equivalent Crack must be defined to determine a retirement time using the Crack Growth Retirement method.
The size of the BEC should bound the life reducing effect of damage that could occur as a result of manufacturing, maintenance, or the service environment. The size may be established by analytical back calculations from coupon or service fatigue life data accounting for material variability effects in the data. In any case, there should be no Page C - 113
AC 29-2C, Chg 4 12/2/2011 probable damage from any source that would lead to failure of the part in less time that it would take the BEC to reach critical size. Each applicant must justify the BEC sizes used in the analysis; however, there has been some limited experience that indicates that the following BEC sizes could be appropriate.
(1) 0.015 inch or 0.380 mm radius semicircular surface crack for precision-machined mechanical parts
(2) 0.050 inch or 1.270 mm radius quarter-circular corner crack in fastener holes for typical aluminum airframe structure (F) Initial Inspectable Crack. The size and shape of the initial inspectable crack (aDET) must be established when the Crack Growth Inspection approach is used.
The inspection interval is based on the time for the initial inspectable crack to grow to a size (aCRIT) that would result in catastrophic failure of the rotorcraft if limit loads were applied. The initial inspectable crack is a function of the inspection method that is used to detect it. Regardless of the inspection method, the probability of detecting this size crack should be high and it should be substantiated.
(6) Inspectability and Inspection Methods. This section provides guidance on selecting and substantiating damage detection methodology for use with the methods of paragraphs f.(8) (Inspection Intervals) and f.(10) (Approved Equivalent Means). The methods of paragraph f.(8) can result in a mandated inspection program that must be included in the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness in accordance with § 29.1529 of the regulatory requirements. Qualified personnel must conduct these inspections at the specified interval using the approved method or methods. Additionally, § 29.571 allows that substantiation may be accomplished by “Approved Equivalent Means,” which is discussed in paragraph f.(10).
These Approved Equivalent Means may include actions that detect damage or flaws indirectly, and are substantiated using the methods of paragraph f.(8). These actions should be shown to be reliable and systematically conducted by knowledgeable personnel. The following are considerations for establishing inspections, inspection methods, or indirect damage detection.
(i) Inspectability. The ease of conducting an inspection should be a design goal for principal structural elements. Design features such as open construction, access panels or ports, or other easy access to fatigue critical areas for needed inspections should be considered. A design that requires disassembly in order to conduct a required inspection, other than during a scheduled maintenance disassembly, should be avoided.
(ii) The specific inspection methods that are used to accomplish fatigue substantiation should be:
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(A) Compatible with the threats identified in the threat assessment, paragraph f.(5), and provide a high probability of detection in the threat assessment and their development, under the operational loads and environment.
(B) Consistent with the capabilities, facilities, and resources of the potential operators of the helicopter. The need to conduct complex or difficult field- level inspections should be avoided, especially when the projected usage of the helicopter may include extended periods of operation in remote areas.
(C) Developed and substantiated for each specific application by means of a full-scale test program, or by experience with similar methods in similar applications.
(D) Included in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness in accordance with § 29.1529 as required by § 29.571(g).
(iii) Detectable Damage Size Assessment.
(A) In the case where the substantiation is predicated on the detection of a specific flaw or crack size, an assessment should be conducted to assure that the selected inspection method would be highly reliable in detecting that size of damage in service. This assessment may be based on the known capability of currently available inspection methods and equipment, provided that this capability is verified by a full - scale test program or by experience with the method in service for similar structure and damage.
(B) If the current capability of a specific inspection method is in question, or if the capability of a specific method needs to be extended to a smaller damage size, then a systematic assessment and substantiation of the method for the intended purpose is appropriate. This assessment could include the determination of the Probability of Detection (POD) as a function of damage size and should consider the capabilities of the potential operators of the helicopter and the environment in which the inspections will be conducted.
(iv) Indirect Detection of Damage. Several damage detection procedures are available that could be used as “Approved Equivalent Means” to support substantiation of a structure [reference paragraph f.(10)]. These procedures, if systematically required and conducted by knowledgeable personnel, can be used in conjunction with the methods presented in paragraph f.(8) to achieve the substantiation. Examples of this type of substantiation are:
(A) In-flight damage detectable by vibration, noise, or observing a bladeout-of-track tip path plane. Consideration should be given to the background levels of noise and vibration, as well as whether the indication is of a different character (more detectable) rather than just a change in level (less detectable).
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AC 29-2C, Chg 4 12/2/2011 (B) Damage that is obvious in a preflight check or routine visual examination. This could include obvious flaws or cracking, but also could include structure that is found to be loose, broken, or soft when deflected by hand. Other obvious damage detection could include fluid leaks, missing fasteners, structure bent or out of alignment, or jamming of mechanical parts.
(C) Damage that is indicated following flight completion. Spectrographic oil analysis would be an example.
(D) Damage detection by automated means. This includes crack detection by foil, fiber, or wire break, load monitoring (to detect a change in internal load distribution), acoustic emission monitoring, or other on-board sensors that meet the goals of damage detectability and reliability.
(7) Retirement Times. Each of the four methods below provides a means to establish a retirement time for each PSE. The determination of the fatigue tolerance characteristics should include an assessment using the conventional Safe-Life methodology. In addition, this serves as a baseline for comparison to retirement times determined with flaws and defects included, and should be used as the structure’s retirement time if it is the lowest calculated time.
(i) The conventional Safe-Life methodology accounts for damage induced by fatigue loading but does not account for flaws and defects due to manufacturing and in-service conditions. If the retirement time is established using this method, then the damage identified in paragraph f.(5) (as required by § 29.571(d)(iii)) must be addressed by inspections or other equivalent means. Information to guide a fatigue evaluation based on a conventional Safe-Life approach is provided in detail in AC 27-1B MG 11.
The method consists of:
(A) Establishing mean fatigue curves (e.g., stress-life or strain-life) based on crack initiation in constant-amplitude or spectrum testing of as-manufactured structure;
(B) Establishing working fatigue curves with strength and life margins; and (C) Conducting a cumulative damage working life calculation using known flight loads and estimated usage.
(ii) A Safe-Life retirement time substantiation with BDF provides a safe period of operation of a structure with probable flaws that may remain in place without detection for that period. Barely detectable flaws are intended to conservatively represent a worst-case of undetectable flaws. The substantiation is accomplished by testing and analysis employing conventional Safe-Life methodology except that an intrinsic and discrete critical flaw in critical locations on the structure is considered. It should be noted that this method, since it is a Safe-Life (crack initiation) method, is not appropriate for use when the flaw being considered is already a crack.
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(A) The types, sizes, and locations of flaws to be considered are determined by the threat assessment (paragraph f.(5)). These flaws may be represented by “equivalent flaws” if it is demonstrated that they have the same or a more severe strength-reducing effect than the corresponding representative flaws.
(B) The mean fatigue strength of the structure with flaws may be determined by one of the following three methods:
(1) Testing a full-scale structure with flaws:
(i) Representative flaws as determined by the threat assessment, or equivalent flaws if substantiated, are imposed at the critical locations on the structure where flaws are likely to occur.
(ii) S-N or spectrum safe-life fatigue testing is conducted; see paragraph e of AC 27-1B MG 11.
(iii) A mean S-N curve with flaws is derived directly from this data.
(2) As-manufactured structure strength modified by the effect of flaws.
(i) A mean strength for as-manufactured structure (without flaws) can be determined using full-scale S-N or spectrum safe-life fatigue testing.
(ii) The effect of flaws may be determined by analysis, by similarity to components where the effect of the flaws has previously been determined, or by a specimen test program incorporating the pertinent features of the full -scale component.
Consideration should be given to the material form, geometric features, surface finish, and steady and vibratory load levels, in combination with flaws representative of those identified in the threat assessment.
(iii) The effect of the flaws is combined with the fatigue result determined on the as-manufactured structure without flaws.
(3) Analytical mean strength modified by the effect of flaws:
(i) A mean strength for as-manufactured structure (without flaws) can be determined analytically, provided that correlation with a similar design can be accomplished, or if additional conservatism is included in the working curve reductions employed in paragraph f.(7)(ii)(C).
(ii) The effect of flaws may be determined by analysis, by similarity to components where the effect of the flaws has previously been determined, or by a specimen test program incorporating the pertinent features of the full -scale component.
Consideration should be given to the material form, geometric features, surface finish, Page C - 117