AC 29-2C, Chg 3 9/30/2008
SUBPART C - STRENGTH REQUIREMENTS
FATIGUE EVALUATION AC 29.571. § 29.571 FATIGUE EVALUATION OF FLIGHT STRUCTURE.
a. Explanation. An evaluation is required to assure structural reliability of the rotorcraft in flight. This evaluation may take the form of either tests or analysis. During the certification process, fatigue testing is more effective than analysis alone in identifying and preventing cracking that may occur during service. Analysis used for substantiation should be validated by tests. AC 27 MG 11 contains background information and acceptable means of compliance with the requirements pertaining to the safe life methodology. A safe life may be assigned or the structure may be determined to be fail safe or a combination of these may be used. AC 29 MG 11 contains background information and acceptable means of compliance with the requirements pertaining to fatigue and flaw tolerance.
b. Procedures.
(1) The fatigue evaluation requires consideration of the following factors:
(i) Identification of the structure/components to be considered.
(ii) Th e stress during operating conditions.
(iii) The operating spectrum or frequency of occurrence.
(iv) Fatigue strength, and/or fatigue crack propagation characteristics, residual s trength of the cracked structure.
(2) Since the design limits, e.g., rotor RPM (maximum and minimum), airspeed, and blade angles (thrust, weight, etc.) affect the fatigue life of the rotor system, it is necessary that flight conditions be conducted at limits that are appropriate for the particular rotorcraft and at the correct combination of these limits. It will be the responsibility of flight test personnel to determine that the flight strain program includes conditions of flight at the various combinations of rotor RPM, airspeed, thrust, etc., that will be representative of the limits used in service. The flight test personnel should assure that the severity of the maneuvers to be investigated is such that actual service use will not be more severe. Flight test verification may be achieved through:
(i) Flying a representative set of maneuvers with the applicant’s pilot in the test aircraft at noncritical combinations of weight, CG, and speed. (An FAA/AUTHORITY letter for specific test authorization would ordinarily be required.)
(ii) Flying a representative set of maneuvers with the applicant’s pilot in a similar (certified) model to assess and agree upon the required maneuvers, control C – 94
deflections, and aircraft rates. The required maneuvers or conditions will be specified in the flight strain program plan.
(iii) Flying a chase aircraft which has a flight envelope appropriate to allow visual confirmation of the proposed and programmed flight maneuvers.
(iv) Observation of telemetered flight data to assure desired control deflections, rates, and aircraft attitudes.
(v) Some combinations of items b(2)(i) through b(2)(iv) above.
(3) Assessing the operation spectrum and the flight loads or strain measurement program will involve airframe, propulsion, and flight test personnel.
(4) Variation in the operating or loading spectrum among models, and variations in the spectrum for a particular model rotorcraft, should be evaluated.
Figure AC 27 MG 11-7 contains typical flight load measurement program conditions to be investigated. An example of a twin turbine spectrum is presented in Figure AC 27 MG 11-9. The tables should be used only as a guide and should be modified as necessary for each particular rotorcraft design.
(5) The difference in loading spectrum for different models that may be anticipated is illustrated by comparing the percentage of time assigned to level flight conditions, specifically 0.8 VH to 1.0 VH for three different rotorcraft designs where VH is the maximum airspeed at maximum continuous power in level flight. The first column applies to a single-piston-engine powered small rotorcraft used in utility operations. The second was obtained from data for a single-turbine-engine powered seven-place small business and utility rotorcraft. The third was obtained from data for a twin-engine-powered 13 passenger transport rotorcraft. It should be noted that the level flight percentage of occurrences shown in the table below for the turbine utility business and turbine transport rotorcraft are examples of a particular design. The high percentage of time shown in this flight regime could be unconservative for some designs, especially if the stresses under these design conditions produce an infinite fatigue life for the particular component. The fatigue spectrum percentage of occurrences may be modified according to the intended operation usage of the rotorcraft. However, a conservative application should be considered. This variation illustrates the "tailoring" of the loading spectrum for the type of rotorcraft and the anticipated usage.
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AC 29-2C, Chg 1 2/12/2003 FIGURE AC 29.571-1 Comparison Percent of Time in Level Flight Turbine Piston Utility Twin Turbine Utility Business Transport
0.8 VNE 25% 0.8 VH 16% 0.8 VH 15%
1.0 VH 15% 0.9 VH 21% 0.9 VH 20%
1.0 VNE 3% 1.0 VH 24% 1.0 VH 38% Total 43% 61% 73% This variation illustrates the “tailoring” of the loading spectrum for the type of rotorcraft and the anticipated usage.
(6) External cargo operations are a unique and demanding operation. A “logging” operator may use 50 maximum power applications per flight hour to move logs from a cutting site to a hauling site. Power is used to accelerate, decelerate, or hover prior to load release. Lifting loads over an obstruction or natural barrier is another example of very frequent high power applications for takeoff and for hovering over the release area. Similar types of operations require flight loads data to assess the effects on fatigue critical components.
(7) Frequently the applicant may request approval of a gross weight for an external cargo configuration that exceeds the standard configuration gross weight. The external cargo VNE is typically significantly lower than the standard configuration VNE possibly due to adverse effects on flight loads at the increased weight.
(8) The impact of the external cargo operation on standard configuration limits should be assessed to determine whether or not the component service lives will be affected. The assessment may be done by calculating an “external cargo configuration” service life for each critical component. The lowest service life obtained from standard configuration flight loads data and loading spectrum, or from external cargo configuration flight loads data and loading spectrum is generally the approved service life. This procedure avoids prorating the operating time between the two types of operations. This procedure is necessary since the regulatory maintenance and operating rules do not require recording time in service for the different types of operations.
(9) The applicant should plan to conduct a flight loads survey program for both a standard configuration and an external cargo configuration, if appropriate. This procedure will avoid delays associated with reinstallation and calibration of equipment.
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AC 29.571A. §29.571 (Amendment 29-28) FATIGUE EVALUATION OF STRUCTURE.
a. Explanation. Amendment 29- 28 adds a requirement to substantiate tolerance to flaws during the fatigue evaluation of structure. A flaw tolerant safe- life evaluation or a fail-safe (residual strength after flaw growth) evaluation is required by § 29.571(b) unless “the applicant establishes that these fatigue flaw tolerant methods for a particular structure cannot be achieved within the limitations of geometry, inspectability, and good design practices.”
b. Procedures.
(1) AC 29 MG 11 provides acceptable general procedures for complying with Amendment 29-28.
(2) Specific rotorcraft drive system gear fatigue evaluation procedures, which supplement AC 29 MG 11, follow:
(i) Fatigue test evidence is necessary for the fatigue evaluation of gears.
The test evidence should be provided by rotating tests of complete gearbox specimens operating under power. The tests provide the basis for analysis leading to the establishment of safe-life.
(ii) The tests are conducted specifically for the purpose of gear tooth evaluation, and components subjected to the tests do not have to be considered serviceable on completion of the test. Excessive wear on bearings and shafts and marking (including spalling) of bearings and gear teeth are acceptable provided no fatigue damage is evident on the gear teeth. However, fatigue damage other than tooth fatigue should be considered for test validity and the integrity of the affected part confirmed as necessary.
(iii) The test conditions (torque versus number of cycles) should permit the setting of mean strength curve(s) to be associated with each primary gear in the drive train. The minimum test condition should encompass those power levels for which repeated application in service is expected under normal conditions. The S -n curve(s), for the material and type of gear, should be reduced by a factor of safety to take into account material and manufacturing variability. The factored curve will then be used in conjunction with the flight power spectrum to determine a life (limited or unlimited) for the gears in the primary drive system.
(iv) Special procedures, which do not affect fatigue evaluation of the gear teeth, may be allowed to facilitate completion of the test provided they have been justified and they do not affect life determination. These include periodic interruption for inspection, replacement of non-critical parts and the use of special lubricants, special cooling systems, and methods to prevent unrepresentative deflections at the test torque levels.
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AC 29-2C, Chg 4 12/2/2011
(v) From evidence in relation to the strength of steel gears of conventional design, it is accepted that adequate fatigue strength can be demonstrated by the use of the above safety factor of 1.4 for a single test, 1.35 for two tests, 1.32 for three tests, and 1.3 for four or more tests. Where several tests are to be conducted, specimens should be selected from different manufacturing batches if practicable.
(vi) Demonstration of infinite life for gear teeth will normally require tests of a minimum of 107 cycles duration at factored power levels. Use of shorter duration tests should be justified.
AC 29.571B. § 29.571 (Amendment 29-55) FATIGUE TOLERANCE EVALUATION OF METALLIC STRUCTURE.
a. Purpose. This advisory material provides an acceptable means of compliance with the provisions of § 29.571 Amendment 29-55 of the FAA regulations dealing with the fatigue tolerance evaluation of transport category rotorcraft metallic structure. This guidance applies to conventional metallic materials. (Corresponding guidance for composite structure can be found in AC 29–2C MG 8, supplemented by AC 20-107B.
Note: once § 29.573 is effective, AC 29.573 will be the current guidance for composite structures.) The fatigue evaluation procedures outlined in this advisory material are for guidance purposes only and are neither mandatory nor regulatory in nature. Although a uniform approach to fatigue tolerance evaluation is desirable, it is recognized that in such a complex area, new design features and methods of fabrication, new approaches to fatigue tolerance evaluation, and new configurations may require variations and deviations from the procedures described herein. It should be noted that § 29.571 requires that the methodology used by the applicant be approved by the FAA to assure compliance with the regulatory requirements.
b. Special Considerations. The unique performance capabilities of rotorcraft and their typical operational environment make fatigue tolerance evaluations both complex and critically important. Due to the many rotating elements inherent in their design, rotorcraft structures are potentially subject to damaging cyclic stresses in practically every regime of flight. The complexity of the fatigue loading is compounded by the fact that rotorcraft are highly maneuverable and are utilized for many widely varying roles.
Corrosion and other environmental damages are not uncommon in rotorcraft operations; neither are inadvertent damages from maintenance that is typically frequent and intensive. For these reasons, special attention should be focused on the fatigue tolerance evaluation of rotorcraft structure.
c. Background.
(1) Fatigue of rotorcraft dynamic components was first addressed in the 1950’s by means of a Safe-Life methodology. The application of this methodology, as Page C - 98
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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AC 29-2C, Chg 4 12/2/2011
(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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AC 29-2C, Chg 4 12/2/2011
(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.
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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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