..
Primacy Load Failure
CRACK GROWIB ~ Secondacy Path Failure Limit 12/2/2011 AC 29-2C, Chg 4 time for the structure being inspected or the rotorcraft design life if the structure has no declared retirement time. Based on the above, (A) Inspection Interval = Lr/N [For N refer to paragraph
f.(8)(ii)(C)].
(B) Limit of validity = NF (i.e., repetitive inspection time would not be valid for operation beyond NF).
Figure AC 29.571B-6: Multiple Load Path Structure Analytical Evaluation to Support Inspection for a Failed Load Path.
(ii) Evaluation by Test. Figure AC 29.571B -7 illustrates some key points if an inspection for a complete load path failure is to be developed based on testing. The inspection interval is based on the test demonstrated residual life (L r) subsequent to load path failure. Because the residual life decreases with the time accumulated prior to a load path failure, there will be a limit of validity to the L r and it will be dependent on the time at which a load path failure is simulated, (N D).
(A) The test article should consist of as-manufactured production parts. Representative “well” condition loading should be applied for some Page C - 125
AC 29-2C, Chg 4 12/2/2011 predetermined period of time, (ND). It is recommended that the “well” condition loading be of sufficient duration so that ND/LSF is not less than the retirement time minus one inspection interval for the structure being inspected or the rotorcraft design life if the structure has no declared retirement time. At the end of this period, the load path that is to be inspected for complete failure should be disabled (e.g., saw cutting, attachment(s) removal, member removal) to simulate its failure. The test should then be restarted with a representative “failed” condition loading. (Note that the external loads may be the same as for the “well” condition if the member failure simulation results in the correct “failed” condition internal load redistribution.) The test should continue until the desired residual life has been achieved or to the time at which the secondary load path can no longer support limit loads without failure, whichever is less, (N0).
(B) In developing the test spectrum, consideration should be given to proper use of representative loads, truncation of non-damaging loads, inclusion of ground-air-ground cycles, clipping of high magnitude loads, and load sequence.
(C) Based on the above,
(a) Demonstrated residual life = Lr = N0-ND.
(b) Repetitive inspection time = Lr/N [For N refer to paragraph f.(8)(ii)(C)].
(c) Limit of validity = ND/LSF.
(d) LSF = 2, Life safety factor.
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Primary Load Path Intentionally Disabled ~ u ''Failed" Condition Loading < ~ U "Well" Condition --------------------------.
~ Loading IN TEST Secondary Load Path Not Capable of Limit Load 12/2/2011 AC 29-2C, Chg 4 Figure AC 29.571B-7: Multiple Load Path Structure Evaluation by Test to Support Inspection for a Failed Load Path.
(4) Inspect for Less Than a Load Path Failure. Inspection for less than a load path failure may require special non-destructive Inspection (NDI) procedures but will result in longer inspection intervals. Figure AC 29.571B -8 illustrates how inspection intervals could be established on the basis of crack growth and residual strength evaluation.
(i) In this case, the inspection interval is based on the life of the secondary load path (Lr) subsequent to primary load path failure at NF plus the time (LP) for a detectable crack (aDET) in the primary load path to grow to critical size under in-service loads. The determination of L r is the same as discussed in paragraph
f.(8)(ii)(B)(3)(i).
(ii) Based on the above, (A) Repetitive Inspection = (LP + Lr)/N [For N refer to paragraph
f.(8)(ii)(C)].
(B) Limit of validity = NF.
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AC 29-2C, Chg 4 12/2/2011 NF Lr
CRACK SIZE ai aDET LP Crack Growth in Primary Load Path Crack Growth in Secondary Load Path after Primary Load Path Failure Failure of Secondary Load Path at Limit Load Failure of Primary Load Path Under Inservice Loads Crack Growth in Secondary Load Path with Primary Load Path Intact Dai
CRACK GROWTH LIFE Figure AC 29.571B-8: Multiple Load Path Structure Analytical Evaluation to Support Inspection for Less than a Failed Load Path.
(C) Safety Factors.
(1) In determining the factor of N to be used for determining the inspection time, consideration should be given to the crack growth data used (e.g., top of scatter data versus average data, number of specimens used to generate data, etc.)
and the capability of the inspection procedure.
(2) The minimum suggested N value should be N=2 in the case where the conservative top-of-scatter crack growth data are used in the crack growth analysis, or N=4 when the average crack growth data are used in the crack growth analysis, or when the crack growth life is obtained from the crack growth test of one specimen (for two or more full scale specimens, N=3 of the shortest crack growth life can be used).
(iii) Safe-Life Inspection for a Failed Element.
(A) A Safe-Life Inspection substantiation for a Failed Load Path provides a safe interval of operation between repetitive inspections for the failed load path. The substantiation is accomplished by testing and analysis employing conventional Safe-Life methodology except that the configuration of the structure substantiated is with the Page C - 128
critical load path inoperative and appropriate flaws imposed on the remainder of the structure, as determined by the threat assessment.
(B) The method described in paragraph f.(8)(i) can be employed for this case with the following differences:
(1) The principal “flaw” considered is failure or loss of the most critical load path. The load path failure can be the result of fatigue cracking, static failure, or a fractured or missing fastener, as determined by the threat assessment, paragraph f.(5).
(2) The remainder of the structure may be representative of normal manufacturing quality unless the threat assessment indicates that larger damage should exist.
(3) The mean strength for the substantiation should be based on the number of cycles from the first load path failure to the first initiation of cracking at any other point in the remaining structure. Any applied load changes or load distribution changes that occur as a consequence of the load path failure should also be included (bending due to increased deflection, for example).
(4) When the remaining structure may have some pre-existing fatigue damage at the time the first load path fails (due to both load paths being highly loaded, for example), this should be factored into the analysis.
(5) The remaining structure after first load path failure must be shown to have limit load capability, considered as the ultimate loading, except in some cases where no retirement life is provided and fatigue damage is expected (see paragraph
f.(10).
(6) The inspection conducted is for the failed or missing load path.
(9) Retirement Time and Inspection Interval Schedules.
(i) Based on the evaluations required by § 29.571, inspections, retirement times, combinations thereof, or other procedures have been established as necessary to avoid catastrophic failure. These inspections, retirement times, or approved equivalent means must be included in the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness (ICA) as required by § 29.1529 and Appendix A29.4 of the regulatory requirements. These inspections, retirement times, or a combination of both are normally stated in hours time-in-service, but may be stated in other terms, such as engine starts, landings, external lifts, etc.
(ii) The design service life should be specified in the fatigue evaluation methodology that must be approved by the FAA. In any case, routine inspections for wear, fretting, corrosion, cracking, and service damage are appropriate. These routine inspections should be noted in the ICAs (maintenance manual) but are not required to Page C - 129
AC 29-2C, Chg 4 12/2/2011 be contained within the ALS of the ICAs unless they are structural inspection intervals or related structural inspection procedures approved under § 29.571.
(10) Approved Equivalent Means. The requirement includes the possibility that in place of setting retirement times or inspections for damage, some other means may be used. All proposals for ‘equivalent means’ must be submitted to the FAA for approval. Potentially equivalent means to inspection include, but are not limited to:
(i) Indirect detection of damage used to establish a period of safe operation for a structure with the damage present. In this case, the detection is based on the effect of the damage, which may be recognized through:
(A) A warning in flight or during maintenance from a specific feature, sensor, or health monitor, including: oil analysis, chip detector, crack detection wire or foil, health monitoring, fluid leaks or pressure change in a sealed chamber; or by (B) Pilot sensitivity to a change in the rotorcraft’s behavior (such as poor blade tracking, noise generation, vibration generation) provided it is well defined and does not require exceptional piloting skills to recognize these behaviors.
(ii) In all cases, an adequate level of residual strength is demonstrated for the period of operation concerned. Generally, limit load will be considered the minimum residual strength requirement. However, load levels less than the critical limit load conditions may be acceptable for consideration of obvious damage sustained in flight and for the completion of that flight only, provided it allows for continued safe flight and landing.
(iii) Two instances are considered here where it may not be necessary to provide a retirement time in the ALS of the ICAs. However, this does not preclude the investigation of fatigue behavior throughout the life of the rotorcraft or of the part if longer.
(A) When fatigue cracking occurs, or is expected to occur, for a specific PSE while in service, then the first approach allows the PSE to operate until the damage is found. Therefore, the inspection must find the damage prior to loss of ultimate load capability. This approach may not be appropriate for a single load path structure. For such a process to be safe, the behavior of the part and associated parts that influence its fatigue behavior must be substantiated for as long as they remain in service. All potential failure modes throughout the life of the rotorcraft must be identified and shown to be consistent, repeatable and addressed by the inspection program. In order to meet the intent of the new fatigue tolerance requirements, a high probability of ultimate load capability is required throughout the lifetime of the component. Therefore, for cracks or other damage that are allowed or highly likely to exist, ultimate load capability should be substantiated for that damage and any growth that may occur during the subsequent inspection period.
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(B) It may be acceptable that a PSE does not have a specific retirement time when the fatigue tolerance of the part, including any damage not controlled by an acceptable inspection program, has been demonstrated to be in excess of the rotorcraft design life to such an extent that no safety benefit arises from imposing that requirement.
(11) Supplemental Procedures.
(i) The requirement states that if inspections, for any of the damage types identified during the threat assessment, cannot be established within the limitations of geometry, inspectability or good design practice, then supplemental procedures must be established that will minimize the risk of each of these types of damage being present or leading to catastrophic failure. When assessing good design practice, measures such as improved protection against impact, scratches, and corrosion should already have been considered. If the part cannot be redesigned to reduce the acquisition and influence of damage, then supplemental procedures should be introduced.
(ii) Supplemental procedures that should be considered include, but are not limited to:
(A) Specifying shorter than usual calendar inspection intervals to reduce the probability of occurrence and the extent of the damage.
(B) Improving control of maintenance processes associated with the component and damage type, such as by providing specifically designed tooling and requiring additional quality checks after each operation is performed.
(C) Introducing an overhaul program.
(D) Restricting the allowable repair limits for the part.
(E) Modifying the PSE design based on service experience if this shows the original design assumptions to be overly conservative with respect to demonstrating impracticality at certification.
(F) Specifying a conservative inspection interval, if the calculated interval cannot be established and there are no other alternatives.
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AC 29-2C, Chg 4 12/1/2011 AC 29.573. § 29.573 (Amendment 29- 54) DAMAGE TOLERANCE AND
FATIGUE EVALUATION OF COMPOSITE ROTORCRAFT
STRUCTURES
a. Purpose. This advisory material provides an acceptable means of compliance with the provisions of § 29.573, Amendment 29-54, Title 14 of the Code of Federal Regulations (CFR) dealing with the damage tolerance and fatigue evaluation of transport category composite rotorcraft structures. Paragraph f.(6) specifically addresses the advisory guidance applying to damage tolerance and fatigue evaluation as required by § 29.573, Amendment 29-54. Some information contained in AC 29- 2C, MG 8 (Amendment 29-42) is repeated and updated, as appropriate, to preserve the “building block” approach for analyses of composite rotorcraft structure for compliance to § 29.573, Amendment 29-54 . (Supplemental guidance can be found in AC 20- 107, “Composite Aircraft Structure.”) These procedures address the substantiation requirements for composite material system constituents, composite material systems, and composite structures common to rotorcraft. A uniform approach to composite structural substantiation is desirable, but it is recognized that in a continually developing technical area, which has diverse industrial roots both in aerospace and in other industries, variations and deviations from the procedures described here may be necessary. Deviations from this advisory material should be coordinated in advance with the Rotorcraft Directorate.
b. Special Considerations. Since rotorcraft structure is configured uniquely and is inherently subjected to severe cyclic stresses, special consideration is required for the substantiation of all rotorcraft structure, including composites. This special consideration is necessary to ensure that the level of safety intended by the current regulations are attained during the type certification process for all structure with special emphasis on composite structure because of its unique structural characteristics, manufacturing quality and operational considerations, and failure mechanisms.
c. Background.
(1) Historically, rotorcraft have required unique, conservative structural substantiation because of unique configuration effects, unique loading considerations, severe fatigue spectrum effects, and the specialized comprehensive fatigue testi ng required by these effects. Rotorcraft structural static strength substantiation for both metal and composite structure is essentially identical to that for fixed wing structure once basic loads have been determined. However, rotorcraft structural fatigue substantiation is significantly different from fixed wing fatigue substantiation. Since AC 20-107, as developed, applies to both fixed wing aircraft and rotorcraft, it, of necessity, was finalized in a broad generic form. Accordingly, a need to supplement AC 20-107 for rotorcraft was recognized during type certification programs. One significant difference in traditional rotorcraft fatigue substantiation programs and fixed wing fatigue programs is the use of multiple component fatigue tests for rotor craft programs rather than just one full-scale test. Also, constant amplitude, accelerated load tests are typically used rather than spectrum tests because of the high frequency Page C - 132
loads common to rotorcraft operations. These rotorcraft fatigue tests have traditionally involved the generation of stress versus life or cycle (S-N) curves for each critical part (most of which are subjected to the cyclic loading of the main or tail rotor system) using a monotonic (sinusoidal) fatigue spectrum based on maximum and minimum service stress values. Unless configuration differences or flight usage data dictate otherwise, the monotonic fatigue spectrum’s period is typically based on six ground-air-ground (GAG) cycles for each flight hour of operation. The S -N curves for the substantiation of each detailed part are typically generated by plotting a curved line through three data points (see AC 29-2C, AC 29 MG 11, “Fatigue Tolerance Evaluation of Transport Category Rotorcraft Metallic Structure”). The three data points selected are a short specimen life (low-cycle fatigue), an intermediate specimen life and a long specimen life (high-cycle fatigue). Each raw data point is generated by monotonically fatigue testing at least two full-scale parts to failure or run out for each data point on the S -N curve.
The raw data point values are then reduced by an acceptable statistical method to a single value for plotting to ensure proper reliability of the associated S-N curve.
Order 8110.9, “Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopter and Other Power Transmission Systems” and AC 27-1B, AC 27 MG 11, “Fatigue Evaluation of Rotorcraft Structure”, contain comprehensive discussions of the S-N curve generation process. The rotorcraft S -N curve process contrasts sharply with the fixed wing process of using a single full-scale fatigue article (usually an entire wing or airframe, which constitutes a single full-scale assembly data point), generic material or full-scale assembly S-N data (e.g., Metallic Materials Properties Development and Standardization (MMPDS), formerly the MIL-HDBK-5 for metals; Composites Materials Handbook-17 (CMH-17), formerly the MIL-HDBK-17 for composites; or AC 23-13, “Fatigue, Fail-Safe, and Damage Tolerance Evaluation of Metallic Structure for Normal, Utility, Acrobatic, and Commuter Airplanes”, which replaced AFS-120-73-2 for full-scale assemblies), a non-monotonic spectrum, and relatively large scatter factors to verify or determine the design fatigue life of the full-scale airplane.
(2) Additionally, rotorcraft have employed and mass-produced composite designs in primary structure (typically main and tail rotor blades) since the early 1950’s.
This was 10 or more years before composites were type certificated for primary fixed-wing structure in either military or civil aircraft applications (with some notable limited production exceptions, such as the Windecker fixed wing aircraft). In any case, the early 1950 period was well before a clear, detailed understanding of composite structural behavior (especially in the areas of macroscopic and microscopic failure mechanisms and modes) was relatively common and readily available in a usable format for the average engineer working in this field. It also predated the initial issuance of AC 20-107. Currently, much composite design information is proprietary, either to government, industry or both, and many data gathering methods have not been completely standardized. Consequently, a significant variation from laboratory to laboratory in material property value determination methods and results can exist. The early rotor blade designs (as well as current designs) are by nature relatively low strain, tension structure designs. Also, by nature, these designs are not damage or flaw critical. Thus, by circumstance as much as design, early composite rotor blade and other composite rotorcraft designs incorporated an acceptable fatigue tolerance level of Page C - 133
AC 29-2C, Chg 4 12/1/2011 safety. In the 1980’s, more test data, analytical knowledge, and analytical methodology became available to more completely substantiate a composite design. Current 14 CFR parts 27 and 29 contain many sections to be considered in substantiating composite rotorcraft structure. This advisory material provides the current or updated information from AC 29-2C, MG 8, Amendment 29-42 to supplement the general guidance of AC 20-107 and provides compliance guidance for the requirements of § 29.573 Amendment 29-54 for rotorcraft composite structure.
d. Definitions. The following basic definitions are provided as a convenient reading reference. CMH -17, and other sources, contain more complete glossaries of definitions.
(1) A-Basis Allowable. The “A” mechanical property value is the value above which at least 99 percent of the population of values is expected to fall, with a confidence level of 95 percent.
(2) Accidental Damage. Discrete damage, which may occur in service use or in manufacturing due to impacts or collisions, such as dents, scratches, gouges, abrasions, disbonds, splintering, and delaminations.
(3) Active Multiple Load Path. Structure providing two or more load paths that are all loaded during operation to a similar load spectrum.
(4) Allowables. Both A -basis and B-basis values statistically derived and used for a particular composite design.
(5) As-Manufactured. Product or component that has passed the applicable quality control process and has been found to conform to the approved design within the allowable tolerances.
(6) Autoclave. A closed apparatus usually equipped with variable conditions of vacuum, pressure, and temperature. It is used for bonding, compressing or curing materials.
(7) B-Basis Allowable. The “B” mechanical property value is the value above which at least 90 percent of the population of values is expected to fall, with a confidence level of 95 percent.
(8) Balanced Laminate. A composite laminate in which all laminae at angles other than 0° occur only in ± pairs (not necessarily adjacent).
(9) Bond. The adhesion of one surface to another, with or without the use of an adhesive as a bonding agent.
(10) Catastrophic Failure. An event that could prevent continued safe flight and landing.
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(11) Cocure. The process of curing several different materials in a single step.
Examples include the curing of various compatible resin system pre-pregs, using the same cure cycle, to produce hybrid composite structure or the curing of compatible composite materials and structural adhesives, using the same cure cycle, to produce sandwich structure or skins with integrally molded fittings.
(12) Component. A major section of the airframe structure (e.g., wing, fin, body, horizontal stabilizer), which can be tested as a complete unit to qualify the structure.
(13) Coupon. A small test specimen (e.g., usually a flat laminate) for evaluation of basic lamina or laminate properties or properties of generic structural features (e.g., bonded or mechanically fastened joints).
(14) Cure. To change the properties of a thermosetting resin irreversibly by chemical reaction (i.e., condensation, ring closure, or addition). Cure may be accomplished by addition of curing (crosslinking) agents, with or without a catalyst, and with or without heat.
(15) Damage. A generic term for structural anomalies caused by manufacturing (processing, fabrication, assembly or handling) or service usage. Trimming, fastener installation, or foreign object impact are potential sources of damage, along with fatigue and environmental effects.
(16) Damage Tolerance. The attribute of the structure that permits it to retain its required residual strength for a period of use after the structure has sustained a given level of fatigue, corrosion, accidental or discrete source damage.
(17) Damage Tolerant Fail-Safe. The capability of structure remaining after a partial failure to withstand design limit loads without catastrophic failure within an inspection period.
(18) Damage Tolerant Safe Life. Capability of structure with damage present to survive expected repeated loads of variable magnitude without detectable damage growth and to maintain ultimate load capability throughout service life of the rotorcraft.
(19) Delamination. The separation of the layers of material in a laminate.
(20) Design Limit Loads. The maximum loads to be expected in service, as defined by § 29.301(a).
(21) Detail. A non- generic structural element of a more complex structural member (e.g., specific design configured joints, splices, stringers, stringer runouts, or major access holes).
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AC 29-2C, Chg 4 12/1/2011
(22) Disbond. A lack of proper adhesion in a bonded joint. This may be isolated or may cover a majority of the bond area. It may occur at any time in the cure or subsequent life of the bond area and may arise from a wide variety of causes.
(23) Element. A generic part of a more complex structural member (e.g., skin, stringers, shear panels, sandwich panels, joints, or splices).
(24) Environment. External, non- accidental conditions (excluding mechanical loading), separately or in combination, that can be expected in service and which may affect the structure (e.g., temperature, moisture, UV radiation, and fuel).
(25) Fatigue or Environmental Damage. Structural damage related to fatigue or environmental effects such as delaminations, disbonds, splintering, or cracking.
(25) Fiber. A single homogeneous strand of material, essentially one-dimensional in the macro-behavior sense, used as a principal constituent in advanced composites because of its high axial strength and modulus.
(26) Fiber Volume. The volume of fiber present in the composite. This is usually expressed as a percentage volume fraction or weight fraction of the composite.
(28) Fill. The 90° yarns in a fabric, also called the woof or weft.
(29) Glass Transition. The reversible change in an amorphous polymer or in amorphous regions of a partially crystalline polymer from (or to) a viscous or rubbery condition to (or from) a hard and relatively brittle one.
(30) Glass Transition Temperature. The approximate midpoint of the temperature range over which the glass transition takes place.
(31) Hybrid. Any mixture of fiber types (e.g., graphite and glass).
(32) Impregnate. An application of resin onto fibers or fabrics by several processes: hot melt, solution coat, or hand lay-up.
(33) Intrinsic or discrete manufacturing defects. Intrinsic or discrete imperfections or flaws related to manufacturing operations, processing or assembly, such as voids, gaps, porosity, inclusions, fiber dislocation, disbonds, and delaminations.
(34) Lamina. A single ply or layer in a laminate in which all fibers have the same fiber orientation.
(35) Laminate. A product made by bonding together two or more layers or laminae of material or materials.
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