9/30/99 AC 29-2C

Page MG 8 - 1 CHAPTER 3 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT MISCELLANEOUS GUIDANCE (MG) AC 29 MG 8. (Amendment 29-35) SUBSTANTIATION OF COMPOSITE ROTORCRAFT STRUCTURE. a. Reference FAR Sections. §§ 29.305, .307, .571, .603, .605, .609, .610, .611, .613, .629, .923, .927, .931, .1529 and Appendix A. b. Purpose. These substantiation procedures provide a more specialized supplement to the general procedures outlined by AC 20-107A, “Composite Aircraft Structure.” These procedures address 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, some variations and deviations from the procedures described herein will be both necessary and acceptable. Significant deviations from this material should be coordinated in advance with the Rotorcraft Directorate. c. 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 is 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. d. 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 testing 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 for metals is significantly different from fixed wing fatigue substantiation. Since AC 20-107A, 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 20-107A 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 full-scale specimen fatigue tests for rotorcraft programs rather than just one full-scale specimen test. Also, constantAC 29-2C 9/30/99

Page MG 8 - 2 amplitude, accelerated load tests are typically used rather than spectrum tests because of the high frequency 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 (reference Appendix 1 of this AC, “Fatigue Evaluation of Transport Category Rotorcraft Structure (Including Flaw Tolerance)”). 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 specimens (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 Appendix 1 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., MIL-HDBK-5 for metals, MIL-HDBK-17B for composites, or 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) Also, 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 safety. In the 1980’s, more test data, analytical knowledge, and analytical methodology became available to more completely substantiate a composite design. Current9/30/99 AC 29-2C

Page MG 8 - 3 FAR’s 27 and 29 contain many sections (reference paragraph a.) to be considered in substantiating composite rotorcraft structure, but this advisory material is needed to supplement the general guidance of AC 20-107A by providing specific rotorcraft guidance for obtaining consistent compliance with FAR sections applicable to rotorcraft. e. Definitions. The following basic definitions are provided as a convenient reading reference. MIL-HDBK-17, and other sources, contain more complete glossaries of definitions. (1) AUTOCLAVE. A closed apparatus usually equipped with variable conditions of vacuum, pressure and temperature. Used for bonding, compressing or curing materials. (2) ALLOWABLES. Both A- basis and B- basis values statistically derived and used for a particular composite design. (3) BALANCED LAMINATE. A composite laminate in which all laminae at angles other than 0° occur only in ± pairs (not necessarily adjacent). (4) 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 of 95 percent. (5) 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 of 95 percent. (6) BOND. The adhesion of one surface to another, with or without the use of an adhesive as a bonding agent. (7) 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. (8) 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 catalyst, and with or without heat. (9) DELAMINATION. The separation of the layers of material in a laminate. (10) DISBOND. A lack of proper adhesion in a bonded joint. This may be local 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.AC 29-2C 9/30/99

Page MG 8 - 4 (11) 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. (12) 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. (13) FILL. The 90° yarns in a fabric, also called the woof or weft. (14) 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. (15) GLASS TRANSITION TEMPERATURE. The approximate midpoint of the temperature range over which the glass transition takes place. (16) HYBRID. Any mixture of fiber types (i.e., graphite and glass). (17) IMPREGNATE. An application of resin onto fibers or fabrics by several processes: hot melt, solution coat, or hand lay-up. (18) LAMINA. A single ply or layer in a laminate in which all fibers have the same fiber orientation. (19) LAMINATE. A product made by bonding together two or more layers or laminae of material or materials. (20) LOW STRAIN LEVEL. As used herein, is defined as a principal, elastic axial gross strain level, that for a given composite structure provides for no flaw growth and thus provides damage tolerance of the maximum defects allowed during the certification process using the approved design fatigue spectrum. (21) MATERIAL SYSTEM CONSTITUENT. A single constituent (ingredient) chosen for a material system (e.g., a fiber, a resin). (22) MATERIAL SYSTEM. The combination of single constituents chosen (e.g., fiber and resin). (23) MATRIX. The essentially homogeneous material in which the fibers or filaments of a composite are embedded. The resins used in most aircraft structure are thermoset polymers. (24) MAXIMUM STRUCTURAL TEMPERATURE. The temperature of a part, panel or structural element due to service parameters such as incident heat fluxes, temperature, and air flow at the time of occurrence of any critical load case, (i.e., each9/30/99 AC 29-2C

Page MG 8 - 5 critical load case has an associated maximum structural temperature). This term is synonymous with the term “maximum panel temperature.” (25) POROSITY. A condition of trapped pockets of air, gas, or void within a solid materials, usually expressed as a percentage of the total nonsolid volume to the total volume (solid + nonsolid) of a unit quantity of material. (26) PRE-PREG, PREIMPREGNATED. A combination of mat, fabric, nonwoven material, tape, or roving already impregnated with resin, usually partially cured, and ready for manufacturing use in a final product which will involve complete curing. Prepreg is usually drapable, tacky and can be easily handled. (27) RESIN. An organic material with indefinite and usually high molecular weight and no sharp melting point. (28) RESIN CONTENT. The amount of matrix present in a composite either by percent weight or percent volume. (29) SECONDARY BONDING. The joining together, by the process of adhesive bonding, of two or more already-cured composite parts, during which the only chemical or thermal reaction occurring is the curing of the adhesive itself. The joining together of one already-cured composite part to an uncured composite part, through the curing of the resin of the uncured part, is also considered for the purposes of this advisory circular to be a secondary bonding operation. (See COCURING). (30) SHELF LIFE. The length of time a material, substance, product, or reagent can be stored under specified environmental conditions and continue to meet all applicable specification requirements and/or remain suitable for its intended function. (31) STRAIN LEVEL. As used herein, is defined as the principal axial gross strain of a part or component due to the principal load or combinations of loads applied by a critical load case considered in the structural analysis (e.g., tension, bending, bending-tension, etc.). Strain level is generally measured in thousandths of an inch per unit inch of part or microinches/per inch (e.g., .003 in/in equals 3000 microinches/inch). (32) SYMMETRICAL LAMINATE. A composite laminate in which the ply orientation is symmetrical about the laminate midplane. (33) TAPE. Hot melt impregnated fibers forming unidirectional pre-preg. (34) THERMOPLASTIC. A plastic that repeatedly can be softened by heating and hardened by cooling through a temperature range characteristic of the plastic, and when in the softened stage, can be shaped by flow into articles by molding or extrusion.AC 29-2C 9/30/99

Page MG 8 - 6 (35) THERMOSET (OR CHEMSET). A plastic that once set or molded cannot be re-set or remolded because it undergoes a chemical change; (i.e., it is substantially infusible and insoluble after having been cured by heat or other means). (36) WARP. Yarns extended along the length of the fabric (in the 0° direction) and being crossed by the fill yarns (90° fibers). (37) WORK LIFE. The period during which a compound, after mixing with a catalyst, solvent, or other compounding constituents, remains suitable for its intended use. f. RELATED REGULATORY AND GUIDANCE MATERIAL.

Document Title (1) AC 20-95 “Fatigue Evaluation of Rotorcraft Structure” (2) AC 20-107 “Composite Aircraft Structure” (3) AC 21-26 “Quality Control for the Manufacture of Composite Materials” (4) MIL-HDBK-17 “Polymer Matrix Composites Volume 1: Guidelines” g. PROCEDURES FOR SUBSTANTIATION OF ROTORCRAFT COMPOSITE STRUCTURE. The composite structures evaluation has been divided into eight basic regulatory areas to provide focus on relevant regulatory requirements. These eight areas are: (1) fabrication requirements; (2) basic constituent, pre-preg and laminate material acceptance requirements and material property determination requirements; (3) protection of structure; (4) lightning protection; (5) static strength evaluation; (6) damage tolerance and fatigue evaluation; (7) dynamic loading and response evaluation; and (8) special repair and continued airworthiness requirements. Original as well as alternate or substitute material system constituents (e.g., fibers, resins, etc.), material systems (combinations of constituents and adhesives), and composite designs (laminates, cocured assemblies, bonded assemblies, etc.) should be qualified in accordance with the methodology presented in the following paragraphs. Each regulatory area will be addressed in turn. It is important to remember that proper certification of a composite structure is an incremental, building block process which involves phased FAA/AUTHORITY involvement and incremental approval in each of the various areas outlined herein. It is strongly recommended that a FAA/AUTHORITY certification team approach be used for composite structural substantiation. The team should consist of FAA/AUTHORITY engineering, the MIDO inspector(s), the associated Designated Engineering Representatives (DER’s), the associated Designated Manufacturing Inspection Representatives (DMIR’s), and cognizant members of the applicant’s organization. Personnel who are composites specialists (or are otherwise9/30/99 AC 29-2C

Page MG 8 - 7 knowledgeable in the subject) should be primary team member candidates. Once selected, it is recommended that team meetings be held periodically (possibly in conjunction with type boards) during certification to ensure the building block certification process is accomplished as intended. (1) The first area is the fabrication requirementsof § 29.605: (i) The quality control system should be developed considering the critical engineering, manufacturing, and quality requirements and a guidance standard such as AC 21-26, “Quality Control For the Manufacture of Composite Materials.” This ensures that all special engineering, or manufacturing quality instructions for composites are presented, evaluated, documented, and approved, using drawings, process and manufacturing specifications, standards, or other equivalent means. This should be one of the early phases of a composite structure certification program, since this represents a major building block for sequential substantiation work. (ii) Specific allowable defect limits on, for example, fiber waviness, warp defects, fill defects, porosity, hole edge effects, edge defects, resin content, large area debonds, and delaminations, etc., for a particular material system component, laminate design, detailed part, or assembly should be jointly established by engineering, manufacturing, and quality and the associated inspection programs for defect detection created, validated, and approved. Each critical engineering design should consider the worse-case effects of the manufacturing process (maximum waviness, disbonds, delaminations, and other critical defects) allowed by the reliability limitations of the approved inspection program. (iii) If bonds or bond lines such as those typical of rotorcraft rotor blade structure are used, special inspection methods, special fabrication methods or other approved verification methods (e.g., engineering proof tests, reference paragraph g(5)) should be provided to detect and limit disbonds or understrength bonds. (iv) Structurally critical composite construction fabrication process and procurement specifications, for fabricating reproducible and reliable structure, must be provided and FAA/AUTHORITY approved early during the certification process and should, as a minimum, cover the following: (A) Vendor and Qualified Parts List (QPL) Control. Applicants should be able to demonstrate to FAA/AUTHORITY certification team members (both the manufacturing and inspection district office (MIDO) and FAA/AUTHORITY engineering) at any time, that their quality control systems ensure on a continuous basis, that only qualified suppliers provide the basic material constituents or material systems (e.g., pre-pregs) that meet approved material specifications. Recommended guidelines for qualification of alternate material systems and suppliers are contained in MIL-HDBK-17B, Volume I, Section 2.3.2. These methods can also be used, periodically for qualification status renewals of existing material systems and suppliers.AC 29-2C 9/30/99

Page MG 8 - 8 (B) Receiving Inspection and In Process Inspection. Applicants should be able to demonstrate to FAA/AUTHORITY certification team members (both MIDO and engineering), at any time, that their receiving and in-process quality control systems provides products which continuously meet approved material and process specifications. Quality systems should be designed with appropriate checks and balances, such that the necessary statistical reliability and confidence levels for the items being inspected (that are specified by engineering) are continuously maintained. This will require periodic standard inspections and engineering characterization tests on basic constituent and material system samples which should be conducted, as a minimum, on a batch-to-batch basis. The periodic testing necessary to maintain the quality standard should be conducted by the applicants on conformed samples and should be FAA/AUTHORITY-witnessed. (C) Material System Component Storage and Handling. Applicants should be able to demonstrate to FAA/AUTHORITY certification team members (both MIDO and engineering), at any time, that their composite material system (or constituent) storage and handling procedures and specifications provide products which continuously meet approved material and process specifications. Quality systems should be designed with appropriate checks and balances, such that the necessary statistical reliability and confidence levels for the items being inspected (which are specified by engineering) are continuously maintained. This should require, as a minimum, periodic inspections to ensure that proper records are kept on critical parameters (e.g., room temperature “bench” exposure, shelf life, etc.) and that periodic basic constituent and material system characterization tests are conducted, on a batch-to-batch basis. The periodic testing necessary to maintain the quality standard should be conducted by the applicants on conformed samples and should be FAA/AUTHORITY-witnessed. (D) Statistical Validation Level. It is necessary to maintain the minimum required statistical validation level of the quality control system (which should be specified for each critical item or constituent by the approved quality and engineering specifications). The statistical validation level should be defined and approved early in certification. Also, approval and proper usage should be continuously maintained during the entire procurement and manufacturing cycles. (v) Alternate fabrication and process techniques should be approved and should comply with § 29.605. Any alternate techniques should provide at least the same level of quality and safety as the original technique. Any changes should be presented and FAA/AUTHORITY-approved well in advance of the change’s production effectivity. (2) The second area is the basic raw constituent, pre- preg, and laminate material acceptance requirements and material property determination requirements of §§ 29.603 and 29.613. These criteria require application of the critical environmental limits such as temperature, humidity, and exposure to aircraft fluids (such as fuel, oils, and hydraulic fluids), to determine their effect on the performance of each composite9/30/99 AC 29-2C

Page MG 8 - 9 material system. Temperature and humidity effects are commonly considered by coupon and component tests utilizing preconditioned test specimens for each material system selected. Material “A” & “B” basis allowable strength values and other basic material properties (based on MIL-HDBK-17, or equivalent) are typically determined by small scale tests, such as coupon tests, for use in certification work. In the case of composites, determination of these basic constituent and material system properties will almost invariably involve the submittal, acceptance and use of company standards. This is currently necessary because MIL-HDBK-17 has not completed development of “B” basis allowables for inclusion in the handbook. Also, test methods vary somewhat from manufacturer to manufacturer; therefore, individual company results will exhibit some scatter in final material property values. Any company standard which is approved and used should meet or exceed related MIL-HDBK-17 requirements. Material structural acceptance criteria and property determination should, as a minimum, include the following: (i) Property characterization requirements of all material systems (e.g., pre-pregs, adhesives, etc.) and constituents (e.g., fibers, resins, etc.) should be identified, documented, and approved. These requirements, once approved, should be placed in all appropriate procedures and specifications (such as those in paragraph (g)(1) above). (ii) Moisture conditioning of test coupons, parts, subassemblies, or assemblies should be accomplished in accordance with MIL-HDBK-17, other similar approved methods or per FAA/AUTHORITY approved programs. (iii) The maximum and minimum temperatures expected in service (as derived from test measurements, thermal analyses on panels and other parts, experience, or a combination) should be determined and accounted for in static and fatigue strength (including damage tolerance) substantiation programs considering associated humidity induced effects. (iv) The glass transition temperature, Tg, is an important characteristic parameter of amorphous polymers, such as epoxies. It is the temperature below which the polymer behaves like a “glassy” solid and above which it behaves like a “rubbery” solid, i.e., it is the temperature at which there is a very rapid change in physical properties. In actuality, the change from a hard polymeric material to a rubbery material takes place over a narrow temperature range. A composite material will experience a drastic reduction in matrix controlled mechanical material properties when loaded in this temperature range. Since the resin (matrix) is the critical structural constituent in a composite and since Tg exceedance is critical to structural integrity; Tg determination is necessary. The Tg margin methodology of MIL-HDBK-17, Section 2.2.2.1, should be implemented, i.e., the wet glass transition temperature (Tg) should be 50° F higher than the maximum structural temperature (see definition). For any type of resin or adhesive, an acceptable temperature margin using MIL-HDBK-17 techniques (e.g., consideration of limited high temperature excursions) or equivalent methodologies based on tests and/or experience should be established and approved early in the certification process.AC 29-2C 9/30/99

Page MG 8 - 10 In no case should structural strength be degraded below limit load capability on a maximum world wide high temperature day. (v) Local design values should be established by analysis and characterization tests and approved for specific structural configurations (point designs) which include the effects of stress risers (e.g., holes, notches, etc.) and structural discontinuities (e.g., joints, splices, etc.). Proper determination of these values for full-scale design and test should be considered one of the most critical building blocks in substantiating and evaluating a composite structure. These transitional load transfer areas typically produce the highest stresses (and strains) and serve as the nucleation sites for many of the failures (including those due to the relatively low interlaminar strength of composites) that occur in service in a full-scale part or assembly. Small scales tests (such as coupon, element, and subcomponent tests), or equivalent approved testing programs, and analytical techniques should be carefully designed, prepared, and approved to evaluate potential “hot spots” and provide accurate simulations and representations of full-scale article stresses and strains in the critical transition areas. Proper certification work in this area will ensure initial safety and continued airworthiness in full-scale production articles. (vi) The design strain level for each major component and material system should be established and approved such that specified impact damage considerations are defined and properly limited. The effects of the approved strain levels should be established for each composite material using small scale characterization tests and the results should be used to establish or verify the maximum allowable design strain level for each full-scale article. The maximum allowable design strain values selected should also take into account the reliability and confidence levels established for the relevant portions of the quality control system. This methodology is necessary because the amount and size of flaws in the production article may restrict the allowable level of design strain. In a no-flaw-growth design, the maximum specified impact damage and manufacturing flaw size at the most critical location on the part will be a major factor in determining the maximum allowable elastic strain. This design approach is currently selected for nearly all civil and most military applications; since, under normal conditions, only visual inspections are required in the field (unless unusual external damage circumstances such as a hail storm occur) to maintain the initial level of airworthiness (safety). However, many military applications because of their demanding missions, employ scheduled field non-destructive inspection (NDI) maintenance, (such as comparative ultrasonics) to ensure that flaw growth either does not occur, is controlled by approved structural repair, or by replacement of affected parts. To date, civil applications have not been presented that desire a flaw growth, phased NDI approach. Therefore, selection of the full-scale article’s design strain limit based on small scale tests for a no flaw growth design is seen to be extremely important. (vii) Composite and adhesive properties should be determined such that detrimental structural creep does not occur under the sustained loads and environments expected in service. Small scale characterization tests (such as coupon, element, and subcomponent tests) and analysis, which verify and establish the full-scale design9/30/99 AC 29-2C

Page MG 8 - 11 criteria and parameters necessary to ensure that detrimental structural creep in full-scale structure does not occur in service, should be conducted early in certification and should be FAA/AUTHORITY-approved. (viii) Material allowable strength values for full-scale design and testing should be developed using the coupon procedures presented in MIL-HDBK-17 or equivalent. At least three batches of material samples should be used in material allowable strength testing. Company standards should be prepared, evaluated and FAA/AUTHORITY approved early in certification (as part of the building block process), that reflect the material property determination considerations recommended in MIL-HDBK-17 on a equal to or better than basis. (3) The third area is the protection of structureas required by § 29.609. Protection against thermal and humidity effects and other environmental effects (e.g., weathering, abrasion, fretting, hail, ultraviolet radiation, chemical effects, accidental damage, etc.) should be provided, or the structural substantiation should consider the results of those effects for which total protection is impractical. Determination and approval of worst-case or most conservative operating limits, and damage scenarios should be accomplished. Appropriate flammability and fire resistance requirements should also be considered in selecting and protecting composite structure. Usually a hazard analysis is conducted early in certification which identifies the various threats and threat levels for which protection must be provided. This data is then used to construct and submit for approval the methods-of-compliance necessary to provide proper structural protection. (4) The fourth area is the lightning protection requirements of § 29.610. Protection should be provided and substantiated in accordance with analysis and with tests such as those of AC 20-53A and FAA Report DOT/FAA/CT-86/8. For composite structure projects involving rotorcraft certified to earlier certification bases (which do not automatically include the lightning protection requirements of § 29.610), these requirements should be imposed as special conditions. The design should be reviewed early in certification to ensure proper protection is present. The substantiation test program should also be established, reviewed and approved early to ensure proper substantiation. (5) The fifth area is the static strength evaluation requirements of §§ 29.305 and 29.307 for composite structure. Only conservative proven methods of static analysis and failure criteria should be employed. The material stress-strain curve should be clearly established, at least through the ultimate design load, for each composite design. Composite structure should be statistically demonstrated, incrementally, through a program of analysis, coupon tests, minor component ultimate load tests and major component ultimate load tests. The static strength substantiation program should consider all critical loading conditions for all critical structure including residual strength and stiffness requirements after a predetermined length of service, e.g., end of life (EOL) (which takes into account damage and other degradation due toAC 29-2C 9/30/99

Page MG 8 - 12 the service period). Analytical reports and tests should consider all possible failure modes and should include the critical, allowable effects of: (i) Environment (reference paragraphs AC 29 MG 8 g(2) and g(3)). (ii) Service Life (residual limit strength and stiffness demonstration). (iii) Load path loss (fail-safe analysis and limit strength demonstration). (iv) The standard fabrication process and its variability. (v) Impact damage expected during service up to the established threshold of detectability of the field inspection methods to be employed. (vi) Point design and structural discontinuity considerations (e.g., stress risers, joints, etc.). (vii) Unless the ultimate strength of each critical bonded joint can be reliably substantiated in production by NDI techniques (or other equivalent, approved techniques), then limit load capability must be guaranteed by either of the following or a combination thereof: (A) The maximum disbond of each critical bonded joint which will carry limit load is established by test, analysis, or both. Disbonds greater than these values are typically prevented by design features. (B) Each critical bonded joint on each production article should be proof tested to the critical limit load. (viii) For static strength analysis laminae and laminate “A” and “B” basis allowables (determined in accordance with paragraph g(2)) should be used subject to the following conditions unless lower material properties are required by point design considerations (e.g., stress risers, joints, etc.) stiffness requirements (e.g., flutter or vibration margins), fatigue strength (including damage tolerance), or other overriding considerations. (A) When applied loads are distributed through a single load path or single member within an assembly, the failure of which would result in the loss of the structural integrity of the component involved or inability of the rotorcraft structure to carry limit load, the part should be designed, analyzed, and tested using “A” basis allowables. (B) Redundant (fail-safe) structures in which the failure of individual elements would result in applied loads being safely redistributed to other load carrying members without exceeding the limit load capability of the rotorcraft structure may be designed, analyzed, and tested using “B” basis allowables.9/30/99 AC 29-2C

Page MG 8 - 13 (6) The sixth area is the fatigue evaluation requirements of § 29.571. The fatigue evaluation method for the rotorcraft being certified should consider damage tolerance in accordance with AC 20-107A. (i) The safe-life method for composite structure as defined in AC 20-107A is a flaw tolerant safe-life method (e.g., the test specimens consider inherent production flaws and impact damage (reference paragraph (7)(ii)). (ii) Large area disbonds, weak bonds, delaminations, or other defects should be considered in tests or be prevented or be limited by appropriate flaw tolerant special design features and by special manufacturing, maintenance, and inspection procedures. Special attention should be assigned to all pure bond lines (reference paragraph (5)). (iii) Non-fail-safe or partially fail-safe dynamic component structure, which may employ bond lines as the only load path, should be designed to relatively small previously approved values of elastic, ultimate strain for the material system utilized, and should be subjected to full-scale S-N curve testing. Six or more specimens are recommended, as part of the substantiation process. Where practical, flight-by-flight spectrum testing should be used. (iv) All critical safety of flight composite structure must be designed to be flaw (damage) tolerant. Environment degradation and in-service damage critical values are typically included in the flaw tolerance evaluation. All other key factors, such as material selection, manufacturing, and quality assurance controls, and in-service inspection and maintenance, as noted previously, are also to be accounted for. (v) The fail-safe design features of the rotor heads and blade retention systems, other critical primary composite structure, and point design features (e.g., bonded metal-to-composite joints) should be assessed and appropriate inspection programs provided to prevent catastrophic failure from flaw/damage propagation. (vi) The method of generating S-N curves using approved raw data should be demonstrated, evaluated, and approved. (vii) Any limited life items must be identified and placed in the Airworthiness Limitations section of the maintenance manual in accordance with § 29.571. (viii) Load spectra, load truncation methods and all other major aspects of the fatigue evaluation are documented in test proposals and approved. (ix) Flaw growth rates (from initial detectability to the established value for residual strength) must be previously established and closely monitored duringAC 29-2C 9/30/99

Page MG 8 - 14 substantiation. This data should be used to establish special phased inspections and maintenance intervals for critical structure, as required. (7) The seventh major area is the dynamic loading and response requirements of § 29.629 for vibration and resonance frequency determination and separation for aeroelastic stability and stability margin determination for flutter critical flight structure. Critical parts, locations, excitation modes, and separations are to be identified and substantiated. This substantiation should consist of analysis supported by tests and tests which account for repeated loading effects and environment exposure effects on critical properties, such as stiffness, mass, and damping. Initial stiffness, residual stiffness, proper critical frequency design, and structural damping are provided as necessary to prevent vibration, resonance, and flutter problems. (i) All vibration and resonance critical composite structure are identified and properly substantiated. (ii) All flutter-critical composite structure are identified and properly substantiated. This structure must be shown by analysis to be flutter free to 1.1 V NE (or any other critical operating limit, such as V D , for a VSTOL aircraft) with the extent of damage for which residual strength and stiffness are demonstrated. (iii) Where appropriate, crash impact dynamics considerations should be taken into account to ensure proper crash resistance and a proper level of occupant safety for an otherwise survivable impact. (8) The eighth area is the special repair and continued airworthiness requirements of §§ 29.611, 29.1529, and FAR Part 29 Appendix A for composite structures. When repair and continued airworthiness procedures are provided in service documents (including approved sections of the maintenance manual or instructions for continued airworthiness) the resulting repairs and maintenance provisions must be shown to provide structure which continually meets the guidance of paragraphs (1) through (7) of this AC paragraph. All certification based repair and continued airworthiness standards, limits, and inspections must be clearly stated and their provisions and limitations defined and documented to ensure continued airworthiness. In general, no composite repair should be attempted which is out of scope to repairs stated in an approved Structural Repair Manual (SRM) without an engineering design approval by a qualified FAA/AUTHORITY representative (DER or staff engineer). The following minimum criteria should be met in any acceptable composite repair: (i) The repair should be permanent. (ii) The repair should restore the structure to the required strength and stiffness. (iii) The repair should restore all functional requirements.9/30/99 AC 29-2C

Page MG 8 - 15 (iv) The repair should have a negligible weight penalty. (v) The repair should be aerodynamically compatible. (vi) The repair materials should be compatible in all essential aspects with the parent materials. In summary, primary composite structure is an especially critical structure that requires a clearly defined, phased approval (building block) certification process. This process should involve the entire project certification team from a project’s start to its finish so that proper certification is continuously and ultimately achieved. Also, in some special cases, involving new advanced state-of-the-art composite technology, an issue paper may be necessary. However, in the majority of cases (using current composite materials and design philosophy) the applicant’s acknowledged use of this advisory material (as recorded in the type board minutes) should eliminate the need for a separate issue paper.9/30/99 AC 29-2C

Page MG 11 - 1 CHAPTER 3 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT MISCELLANEOUS GUIDANCE (MG)

AC 29 MG 11 FATIGUE EVALUATION OF TRANSPORT CATEGORY ROTORCRAFT STRUCTURE (INCLUDING FLAW TOLERANCE) a. PURPOSE. This advisory material provides an acceptable means of compliance with the provisions of § 29.571 of the Federal Aviation Regulations (FAR) dealing with the fatigue evaluation of transport category rotorcraft structure (AC 20-95, Fatigue Evaluation of Rotorcraft Structure, May 18, 1976, applies to normal category rotorcraft structure and older transport category rotorcraft). The fatigue evaluation procedures outlined in paragraph AC 29 MG 11 are for guidance purposes only and are neither mandatory nor regulatory in nature. Although a uniform approach to fatigue evaluation is desirable, it is recognized that in such a complex problem, new design features and methods of fabrication, new approaches to fatigue evaluation, and new configurations may require variations and deviations from the procedures described herein. It is recommended that major deviations from the procedures be coordinated with the Rotorcraft Standards Staff, ASW-110, to assure national standardization. b. SPECIAL CONSIDERATIONS. The structure of rotorcraft is subject to cyclic stresses in practically every regime of flight. In addition, since rotorcraft are highly maneuverable and capable of forward, rearward, sideward, vertical, and rotational flight, operating limitations due to fatigue are possible in practically all flight situations. Corrosion and other environmental damage are also common in rotorcraft operations. For these reasons, special attention should be focused on the fatigue evaluation of rotorcraft structure.

c. BACKGROUND. (1) During recent years there have been significant state-of-the-art and industry practice developments in the area of structural fatigue and fail-safe strength evaluations of transport category rotorcraft. The advance in the state-of-the-art has resulted from two primary programs: (1) the perfecting of production techniques for composite construction, and (2) the damage tolerant design features required to meet the battle damage requirements of military programs such as the Advanced Attack and Heavy Lift Helicopter (AAH and HLH) programs. (2) Recognizing that advances in state-of-the-art and industry practice warranted changes to the existing fatigue requirements in Part 29, the regulatory requirements of § 29.571 were substantially revised. The revision to § 29.571 requires new guidance material containing compliance provisions related to the changes. Also, paragraph AC 29 MG 11 supplements AC 20-95 for new transport category rotorcraft and provides guidance material in the flaw tolerance area. General guidance materialAC 29-2C 9/30/99

Page MG 11 - 2 for flaw/damage tolerance of composite structures is provided in AC 20-107A, Composite Aircraft Structure, April 25, 1984.

d. INTRODUCTION. (1) Definitions. (i) Fatigue Tolerance. The capability of structure to continue functioning without catastrophic failure after being subjected to fatigue (repeated) loads expected during operation of the rotorcraft. Fatigue tolerance may be achieved by safe-life design, flaw tolerant safe-life design (enhanced safe life), fail-safe design considering flaw growth, or a combination. (A) Safe Life. The capability of pristine structure as shown by tests, or analysis based on tests, not to sustain measurable cracks during the service life of the rotorcraft or before an established replacement time. Special inspection intervals or other special procedures are not usually prescribed for safe-life substantiations. (Routine inspections for wear, fretting, corrosion, crack, and service damage as outlined in § 29.1529 are, of course, appropriate.) (B) Flaw Tolerant (Enhanced) Safe Life. The capability of flawed structure as shown by tests or analysis based on tests to sustain, without measurable flaw growth, the spectrum of operating loads expected during the service life of the rotorcraft or during an established replacement time. Measurable flaw growth means flaw growth beyond an acceptable threshold of detectability. Enhanced safe life may be achieved by designing structure (single element or multiple element) that provides for resistance to crack initiation from manufacturing or service flaws, by material selection, by material processing, by limitation of stress levels, and by geometric design features. The protection of structure against environmental damage and accidental mechanical damage through the use of protective coatings may be used in determining initial flaw types and sizes to be considered.” (C) Fail-Safe Design Considering Flaw Growth. The capability of rotorcraft structure to continue functioning without catastrophic failure after being subjected to fatigue damage, corrosion, intrinsic flaws, or accidental damage expected during fabrication and operation of the rotorcraft. The concept of fail-safety now includes some discrete damage requirements and more explicitly requires consideration of “life-remaining” after flaws occur until the flaws are detected using a prescribed inspection plan or until the part is replaced. Fail-safe designs may include any of the following features: (1) Multiple Load Path . Structure providing two or more separate and distinct paths of structure that will carry limit load after complete failure of one of the members.9/30/99 AC 29-2C

Page MG 11 - 3 (i) Active Multiple Load Path . Structure providing two or more load paths that are all loaded during operation to a similar load spectrum. The use of active multiple load paths requires special attention to fatigue damage in the remaining members after failure of a member. (ii) Passive Multiple Load Path . Structure providing load paths with one or more of the members (or areas of a member) relatively unloaded until failure of the other member or members. (2) Flaw Arrest (Flaw Stopper) Feature . Structure that does not provide completely separate and distinct load paths but does provide features of design such as bonded and/or riveted straps, changes in geometry, or special processing techniques such as rolling or coining to retard or arrest flaw growth. (3) Slow Flaw Growth Feature . Structure (single element or multiple element) that provides for slow flaw growth by material selection, material processing, limitation of stress levels, geometrical design features, or by other methods. NOTE: The detection of the crack or other flaw is an integral part of the flaw growth method. The need to use complicated inspection techniques may not be practical in some cases. See the discussion in paragraph AC 29 MG 11 g(1)(iii)(B). (ii) Flaw. Structural imperfections in excess of type design allowances for “as manufactured” or “pristine” structure. (A) In Metallics. (1) For Fail-Saf e Crack Growth. Corner cracks for holes, semicircular cracks for surfaces, realistic cracks for other locations. (2) For Flaw Tolerance (Enhanced) Safe Life . Gouges, scratches, corrosion, fretting, or wearing likely to occur during fabrication and operation of the rotorcraft. (B) In Nonmetallics. Flaws should be determined in accordance with AC 20-107A guidance. (iii) Limit Design Load. “The maximum loads to be expected in service,” as defined by § 29.301(a), are considered as limit loads for new structure and as ultimate loads for flaw tolerance residual strength demonstration purposes. The residual strength after failure should equal or exceed these loads for flaw tolerance. (iv) Damage Tolerance. No definition is provided in paragraph AC 29 MG 11 for this expression because many definitions have been applied by the civil and military communities. The expression has been used to describe design features, structural systems (including design, manufacturing, and operatingAC 29-2C 9/30/99

Page MG 11 - 4 considerations), and substantiation techniques such as fracture mechanics analyses for metallics. Paragraph AC 29 MG 11 uses expressions for those features of the general damage tolerance concept used to provide and demonstrate fatigue tolerance of rotorcraft structure including consideration of flaws. Tolerance to flaws rather than tolerance to damage is specified in paragraph AC 29 MG 11 since flaws intrinsic to certain manufacturing and fabrication processes are covered in addition to damage from handling, corrosion, etc. (2) Rotorcraft Flaw Tolerance. Flaw tolerant design as substantiated by fail-safe flaw growth or flaw tolerant (enhanced) safe-life means outlined in § 29.571 and paragraph AC 29 MG 11 g is required, unless it entails such complications that an effective flaw tolerant structure cannot be achieved within the limitations of geometry, inspectability, or good design practice. Good design practice includes consideration of component complexity, component weight, methods of production and component cost. Under these circumstances, a design that complies with safe-life criteria should be used. Typical examples of structure that might not be conducive to flaw tolerance design are swashplates, main rotor shafts, push rods, small rotor head components (i.e., devices, bolts, etc.), landing gear, and gearbox internal parts including bearings. In addition, the need for the use of inspection techniques and equipment or highly trained personnel--resources not available (for economic or other reasons) to the small operator or in remote areas of operation--should be carefully considered (reference paragraphs AC 29 MG 11 d(1)(i)(C)(3) and g(1)(iii)(B)). (3) Test Background. Experience with the application of methods of fatigue evaluation indicates that a relevant test background should exist in order to achieve the design objective. Even under the flaw tolerance method discussed in paragraph AC 29 MG 11g, it is the general practice within industry to conduct flaw tolerance tests for design information and guidance purposes. Flaw location and flaw growth data based on test results and service history of similar parts, if available, should also be considered in establishing a recommended inspection program. (4) Manufacturing Considerations. Assurance of structural adequacy also includes manufacturing and fabrication in accordance with design requirements and specifications, quality control to monitor compliance, and effective service inspection procedures. (5) Fatigue Tolerance Considerations. In the fatigue tolerance evaluation, the following items should be considered: (i) Identification of the structure to be considered in each evaluation (a failure mode and effects analysis or similar method should be used). (ii) The stresses and strains (steady and oscillatory) associated with all representative steady and maneuvering operating conditions expected in service.9/30/99 AC 29-2C

Page MG 11 - 5 (iii) The frequency of occurrences of various flight conditions and the corresponding spectrum of loadings and stresses. (iv) The fatigue strength, fatigue crack propagation characteristics of the materials used and of the structure, and the residual strength of the damaged structure. (v) Inspectability, inspection methods, and detectable flaw sizes. (vi) Variability of the measured stresses of paragraph AC 29 MG 11 d(5)(ii), the actual flight condition occurrences of paragraph AC 29 MG 11 d(5)(iii), and the fatigue strength material properties of paragraph AC 29 MG 11 d(5)(iv).

e. FLIGHT STRAIN MEASUREMENT PROGRAM. (1) General. Subsequent to design analysis, in which aircraft loads and associated stresses are derived, the stress level and/or loads are to be verified by a carefully controlled flight strain measurement program. (This guidance is similar to that of AC 20-95.) (2) Instrumentation. (i) The instrumentation system used in the flight strain measurement program should accurately measure and record the critical strains under test conditions associated with normal operation and specific maneuvers. The location and distribution of the strain gages should be based on a rational evaluation of the critical stress areas. This may be accomplished by appropriate analytical means supplemented, when deemed necessary, by strain sensitive coatings or photoelastic methods. The distribution and number of strain gages should define the load spectrum adequately for each part essential to the safe operation of the rotorcraft as identified in § 29.571(a)(1)(i). Other devices such as accelerometers may be used as appropriate. (ii) The corresponding flight parameters (airspeed, rotor RPM, center of gravity accelerations, etc.) should also be recorded simultaneously by appropriate methods. This is necessary to correlate the loads and stresses with the maneuver or operating conditions at which they occurred. (iii) The instrumentation system should be adequately calibrated and checked periodically throughout the flight strain measurement program to ensure consistent and accurate results. (3) Parts to be Strain-Gauged. Fatigue critical portions of the rotor systems, control systems, landing gear, fuselage, and supporting structure for rotors, transmissions, and engine are to be strain-gauged. For rotorcraft of unusual or unique design, special consideration might be necessary to ensure that all the essential parts are evaluated.AC 29-2C 9/30/99

Page MG 11 - 6 (4) Flight Regimes and Conditions to be Investigated. (i) Typical flight and ground conditions to be investigated in the flight strain measurement program are given in Attachment 1 to paragraph AC 29 MG 11. (ii) The determination of flight conditions to be investigated in the flight strain measurement program should be based on the anticipated use of the rotorcraft and, if available, on past service records for similar designs. In any event, the flight conditions considered appropriate for the design and application should be representative of the actual operation in accordance with the rotorcraft flight manual. In the case of multiengined rotorcraft, the flight conditions concerning partial engine-out operation should be considered in addition to complete power-off operation. The flight conditions to be investigated should be submitted in connection with the flight evaluation program. (iii) The severity of the maneuvers investigated during the flight strain survey should be at least as severe as the maximum likely in service. (iv) All flight conditions considered appropriate for the particular design are to be investigated over the complete rotor speed, airspeed, center of gravity, altitude, and weight ranges to determine the most critical stress levels associated with each flight condition. The temperature effects on loads as affected by elastomeric components are to be investigated. To account for data scatter and to determine the stress levels present, a sufficient amount of data points should be obtained at each flight condition. Consideration can be given to the use of scatter factors in determining the sufficiency of data points. In some instances, the critical weight, center of gravity, and altitude ranges for the various maneuvers can be based on past experience with similar design. This procedure is acceptable where adequate flight tests are performed to substantiate such selections. The combinations of flight parameters that produce the most critical stress levels should be used in the fatigue evaluation.

f. FREQUENCY OF LOADING. (1) Types of Operation. (i) The probable types of operation (transport, utility, etc.) for the rotorcraft should be established. The type of operation can have a major influence on the loading environment. In the past, rotorcraft have been substantiated for the most critical general types of operation with some consideration of special, occasional types of operation. To assure that the most critical types of operation are considered, each major rotorcraft structural component should be substantiated for the most critical types of operation as established by the manufacturer. The types of operation shown below should be considered and, if applicable, used in the substantiation:9/30/99 AC 29-2C

Page MG 11 - 7 (A) Long flights to remote sites (low ground-air-ground cycles but high cruise speeds). (B) Typical, general types of operation. (C) Short flights as used in logging operations. (ii) One means is to substantiate for the most severe type of operation; however, this method is not always economically feasible. (iii) A second means is to quantify the influence of mission type on fatigue damage by adding to or replacing hour limitations by flight cycle limitations (if properly defined and easily identifiable by the crew, for example: one landing, one load transportation). A special type of flight hour limitation replacement using factorization of flight hours for multiple types of operations may be feasible if continuing manufacturers’ technical support is provided and documented; i.e., the manufacturer either provides the factorization analyses or checks them on a continuing basis for each rotorcraft. (iv) Where one or more of the above operations are not among the general uses intended for the rotorcraft, the rotorcraft flight manual should state in the limitations section that the intended use of the rotorcraft does not include certain missions or repeated maneuvers (i.e., logging with its high number of takeoffs/landings per hour). A note to this effect should also appear in the rotorcraft airworthiness limitations section of the maintenance manual prepared in accordance with § 29.1529. (v) Should subsequent usage of the rotorcraft encompass a mission for which the original structural substantiation did not account, the effects of this new mission environment on the frequency of loading and structural substantiation should be addressed and where practicable, in the interest of safety, a reassessment made. If this reassessment indicates the necessity for revised retirement times, those new times may be limited to aircraft involved in the added mission provided-- (A) Proper part reidentification is established; (B) A Rotorcraft Flight Manual (RFM) supplement outlining limitations is approved; (C) An airworthiness limitations section supplement is approved; or (D) An appropriate combination of part reidentification, RFM supplement, or airworthiness limitation section supplement is approved. (2) Loading Spectrum. The spectrum allocating percentage of time or frequencies of occurrence to flight conditions or maneuvers is to be based on the expected usage of the rotorcraft. This spectrum is to be such that it is unlikely that actual usage will subject the structure to damage beyond that associated with theAC 29-2C 9/30/99

Page MG 11 - 8 spectrum. Considerations to be included in developing this spectrum should include prior knowledge based on flight history recorder data, design limitations established in compliance with § 29.309, and recommended operating conditions and limitations specified in the rotorcraft flight manual. The distribution of times at various forward flight speeds should reflect not only the relation of these speeds to V NE but also the recommended operating conditions in the rotorcraft flight manual that govern V cor cruise speed. Where possible, it is desirable to conduct the flight strain-gage program by simulating the usage as determined above, with continuous recording of stresses and loads, thus obtaining directly the stress/load spectra for structural elements. g. FATIGUE STRENGTH EVALUATION (INCLUDING ROTORCRAFT FLAW TOLERANCE). (1) General. A means should be established using the conventional safe-life approach or another fatigue tolerant approach to control the airworthiness of principal structural elements identified under § 29.571(a)(1)(i). While the conventional safe-life approach is acceptable under certain circumstances as defined in § 29.571, the enhanced safe-life and fail-safe flaw growth approaches are to be used unless shown to be impractical as stated in § 29.571. A fatigue strength evaluation of structure considering tolerance to flaws is intended to ensure that even when flaws are present due to manufacturing of service operations, the structure will withstand service loads without failure until the flawed parts are replaced or until the flaws (including resulting fatigue cracks) are detected and appropriate action taken. Either of two types of fatigue strength evaluation may be used for flaw tolerance substantiation: flaw tolerant (enhanced) safe-life or fail-safe flaw growth methods. Flaw tolerant (enhanced) safe-life includes the testing and analyses currently associated with safe-life substantiation, plus consideration of flaws. Flaw growth methods include the testing and analyses currently associated with damage tolerance assessment (DTA). Tests are required to substantiate flaw propagation rates and residual strength. Either method or a combination can be used to meet the requirements of § 29.571 for tolerance to flaws. Flaw tolerance evaluation encompasses establishing the components to be designed as flaw tolerant, defining the loading conditions and extent of flaws for which the structure is to be designed, conducting structural tests and analyses to substantiate that the design objectives have been achieved, and establishing replacement times or establishing inspection programs as necessary to assure detection of fatigue damage. On components predominantly loaded by centrifugal force, care should be taken in selecting limit load to assure that it is the maximum expected in service. Design features that should be used in attaining a flaw tolerant structure are: (i) Use of multi-path construction and the provision of crack stoppers to limit the growth of cracks and to provide adequate residual strength. (ii) Selection of materials and stress levels that preclude crack growth from flaws or that provide a controlled slow rate of crack propagation combined with9/30/99 AC 29-2C

Page MG 11 - 9 high residual strength after initiation of cracks. Tests are required to substantiate crack propagation rates. (iii) Design to permit detection of cracks and other flaws, including the use of crack detection systems, in all critical structural elements before cracks can propagate and become dangerous or result in appreciable strength loss and to permit replacement or repair. Inspection means appropriate for flaw tolerant design follow: (A) Routine Inspections. To support routine inspection programs, blind areas should be avoided, where practical. Access panels and openings should be considered early in design. (B) Special Inspections. These inspections will generally result from test results as well as the geometry of the design. Care should be given to special inspection techniques to be used in the field. Inspection techniques requiring facilities and resources beyond the capability of the small operator or not generally available in remote-area operations traditionally associated with rotorcraft operations should not be specified for field inspections. Conservative sizes for detectable cracks or other flaws should be used. Sufficient interval inspections should be provided to detect cracks before they grow from a detectable size to a size that reduces the remaining strength below design limit strength. If special inspection techniques not commonly available cannot be avoided, then use of enhanced safe-life substantiation techniques should be used. (C) Pressurized Chambers. This design feature may be used to detect cracks that cause a chamber to lose its pressure (either positive or negative). The loss of pressure can be indicated by gages, or dye may be used if it is shown to be a dependable indicator. Care should be taken in the design of pressurized chambers and their indicating systems to assure dependability. Undependable systems can cause an inordinate number of false indications and maintenance problems. (D) Vibration Generation. This characteristic should be considered both from the aspect of vibrations giving indications of a failure and from the aspect of the increased fatigue loading resulting from the vibrations. (E) Noise Generation. If initial failure will result in a clear and unmistakable noise that is sufficiently continuous and loud, this characteristic can be used in achieving flaw tolerance without additional special inspections. (F) Crack Detection Wire, Foil, etc. Detection wire may be used in areas that are sufficiently well defined so that the wire can be properly located. This technique is appropriate in areas otherwise difficult to inspect. The potential for false readings and possible maintenance problems should be considered. (G) Health Monitoring. Techniques such as vibration sensing and analysis or real time oil analysis can be used to provide information on establishing inspection time.AC 29-2C 9/30/99

Page MG 11 - 10 (iv) Use of multiple element structures may be provided so that damage or failure occurring in one element of the member will be confined to that element and the remaining structure will still possess adequate load-carrying ability until the failed element is discovered by inspection. (v) Provisions to limit the probability of concurrent multiple damage, particularly after long service, should be provided. These provisions should ensure adequate independence of each failure mode of multi-path constructions. The use of full-scale fatigue test articles are recommended in this evaluation. Examples of concurrent multiple damage to be avoided are: (A) Simultaneous failure or partial failure of multiple path discrete elements working at similar stress levels. (B) Failures or partial failures, in adjacent areas, due to redistribution of loading following a failure of a single element. (2) Identification of Principal Structural Elements. Principal structural elements are those that contribute significantly to carrying flight and ground loads and whose failure could result in catastrophic failure of the rotorcraft. Typical examples of such elements are: (i) Rotor blades and attachment fittings. (ii) Rotor heads, including hubs, hinges, and some main rotor dampers. (iii) Control system components subject to repeated loading, including control rods, servo structure, and swashplates. (iv) Rotor supporting structure (lift path from airframe to rotor head). (v) Fuselage, including stabilizers and auxiliary lifting surfaces. (vi) Main fixed or retractable landing gear and fuselage attachment structure. (3) Identification of Locations Within Principal Structural Elements to be Evaluated. The locations of damage to structure for damage tolerance evaluation can be determined by analysis or by fatigue test on complete structures or subcomponents. However, tests will be necessary when the basis for analytical prediction is not reliable, such as for complex components. If less than the complete structure is tested, care should be taken to ensure that the internal loads and boundary conditions are valid. (i) The following should be considered:9/30/99 AC 29-2C

Page MG 11 - 11 (A) Strain gage data on undamaged structure to establish points of high stress concentration as well as the magnitude of the concentration; (B) Locations where analysis shows high stress or low margins of safety; (C) Locations where permanent deformation occurred in static tests; (D) Locations of potential fatigue damage identified by fatigue analysis; (E) Locations where the stresses in adjacent elements will be at a maximum with an element in the location failed; (F) Partial fracture locations in an element where high stress concentrations are present in the residual structure; (G) Locations where detection would be difficult; (H) Design details that service experience of similarly designed components indicates are prone to fatigue or other damage; and (I) Components fabricated from materials of potentially low fracture toughness or high flaw growth rate. (ii) In addition, the areas of probable damage from sources such as a severe corrosive and/or fretting environment, a wear and/or galling environment, or a high maintenance environment should be determined from a review of the design and past service experience. (4) Extent of Flaws. Each particular design should be assessed to establish appropriate damage criteria in relation to inspectability and flaw extension characteristics. In any flaw determination, it is possible to establish the extent of flaws in terms of detectability with the inspection techniques to be used, the associated single element failure or initially detectable flaw size, the residual strength capabilities of the structure, and the likely flaw extension rate (after either an element failure or a partial failure) considering the expected stress redistribution under the repeated loads expected in service and with the expected inspection frequency. Although multiple-element design should be used where practical, and obvious partial failure could be considered to be the extent of the flaw for residual strength assessment, provided a positive determination is made that manufacturing or service flaws will not initiate flaw growth within the life of the part or that the fatigue flaw growth will be detectable by the available inspection techniques at a sufficiently early stage of the flaw growth. In a swashplate or pin containing pressurized chamber, an obvious partial failure might be detectable through the inability of the chamber to maintain pressure after occurrence of the damage. Flaw tolerant (enhanced) safe-life evaluations from flaws should consider flaws to be expected during manufacturing (including handling) and during service. Special coatings against corrosion, flame plating or plasma platingAC 29-2C 9/30/99

Page MG 11 - 12 against fretting corrosion, and energy absorption coating or shielding against damage associated with maintenance may be used in determining the type and extent of flaws to be considered in flaw tolerant (enhanced) safe-life evaluations. The following are typical examples of the type of partial failures that may be considered in the flaw growth fail-safe and enhanced safe-life evaluation. These may or may not be appropriate to the design being considered. These examples have been sources of service difficulty on prior/existing designs: (i) Detectable skin cracks in the trailing edge sections of rotor blades. (ii) Detectable failures of individual straps in “strap packs.” (iii) Detectable skin cracks emanating from the edge of structural openings or cutouts; (iv) Detectable circumferential or longitudinal skin crack in the basic fuselage or tail boom structure; (v) Complete severance of interior frame elements or stiffeners in addition to a detectable crack in the adjacent skin; (vi) Presence of a detectable fatigue failure in at least the tension portion of the spar web or similar element; (vii) Detectable failure of a primary attachment, including blade attachment fittings and control surface hinge and fittings; and (viii) Fretting, corrosion, and galling conditions expected in service. (5) Provisions for Inspection. The designer should strive to ensure adequate inspectability of all structural parts to qualify them under the fail-safe flaw growth provisions. In those cases where blind areas or surfaces exist, suitable design features should be provided to allow inspection techniques (either visual or nondestructive testing, as necessary) to assure adequate residual strength is achieved unless shown to be impractical due to limitations of geometry and good design practice. In addition, the alternate safe-life approach to fatigue tolerance should be implemented if the inspection techniques are shown to be too complicated and impractical. (6) Testing of Principal Structural Elements. The nature and extent of tests on complete structures or on portions of the primary structure will depend upon applicable previous design, construction, tests, and service experience in connection with similar structures. For flaw tolerant safe-life testing, simulated flaws should be as representative as possible of actual gouges, scratches, pitting, or fretting to be expected in manufacture and service. For fail-safe testing considering crack propagation, simulated cracks should be as representative as possible of actual fatigue damage. Where it is not practical to produce actual fatigue cracks, flaws can be9/30/99 AC 29-2C

Page MG 11 - 13 simulated by cuts made with a fine saw, sharp blade, guillotine, or other suitable means. The validity of saw cuts, etc., should be verified by comparison to coupon tests of a cracked specimen of the same material. In those cases where bolt failure, or its equivalent, is to be simulated as part of a possible flaw configuration in joints or fittings, bolts can be removed to provide that part of the simulation. (7) Flaw Tolerance Demonstration (Flaw Tolerant (Enhanced) Safe-Life or Fail-Safe Flaw Growth). (i) It should be determined by analysis, supported by test evidence, that the structure with the extent of damage established for residual strength evaluation can withstand the specified design limit loads (considered as ultimate loads). Flaw tolerant safe-life substantiation provides a safe period of operation of structure with flaws with only routine inspections necessary. Safe crack growth (fracture mechanics) substantiation for fail-safe designs, on the other hand, provides for limited operation after crack initiation from flaws until the cracks can be safely detected. Since flaw tolerant safe life does not provide for detailed crack growth beyond the threshold of detectability, it will tend to apply to life-limited parts, particularly those hard to inspect. Safe crack growth substantiation of fail-safe designs is applicable to readily inspectable parts and may provide for replacement “on condition” rather than at a specified life. (ii) The enhanced safe-life of metallic components will use analyses and testing similar to that of basic safe-life except that “flawed” specimens will be tested rather than “pristine” specimens. (A) In order to determine the mean fatigue strength and the variability in fatigue strength considering flaws, it is necessary to test a number of specimens in establishing stress versus number of cycles (S-N) curves. Both full-scale and coupon specimens may be used to account for the variability in fatigue strength. A reduction factor should be applied to the mean curve in arriving at a working S-N curve. This factor should include consideration of the number of specimens tested, the variability of the fatigue results and, where available, previous test data on the same materials or similar components, as well as service experience. (B) Where new materials or designs are being evaluated, it is recommended that a larger reduction factor be used until additional test data justifying a change are available. The mean and reduced S-N data acceptable to the Administrator on specimens with stress concentration factor, as applicable. A reduced S-N curve and the loading spectrum of paragraph AC 29 MG 11 f(2) should be used in determining replacement times. (C) Figure AC 29 MG 11-1 represents the method of constructing a typical S-N curve from the fatigue test data. Four to six full-scale specimens have commonly been used in past safe-life programs to determine mean S-N curves. The applicant may propose a specific number of specimens that may be evaluated withAC 29-2C 9/30/99

Page MG 11 - 14 respect to the proposed methodology in determining the acceptability of the structural substantiation program.9/30/99 AC 29-2C

Page MG 11 - 15 (D) Additional coupon tests have been used, when necessary, to identify more completely test scatter due to material properties variability. When the agreed-upon number of specimens is tested at various acceptable load levels, a reduction by three mean standard deviations, sigma ( σ), necessary to account for material properties, should be made to establish a reduced S-N curve for determination of retirement times. This reduction may take into account the fact that flawed specimens were tested to preclude a dual penalty situation. Reduction by two mean standard deviations rather than three, may be used if justified by appropriate design features such as multiple elements or unmistakable flaw indications or by material properties that provide benign types of failure modes. (E) “Run-outs” (specimens that do not fail) may be treated in one of three ways: deleted from data (conservative), considered as a failure at maximum test cycles (conservative), or considered in a rational manner to develop a more realistic “ σ .” To further account for variability in fatigue strength not totally accounted for by this reduction in S-N data, the lesser of this reduction, 80 percent of mean, or bottom of scatter, may be used as a limit, where justified, depending on material properties. (iii) The procedures in Paragraph 7a of AC 20-107A provide criteria for substantiating safe flaw growth for composite structure. (iv) Safe crack growth (fracture mechanics) substantiation should show that the damage growth rate under the repeated loads expected in service (between the time at which the damage becomes initially detectable and the time at which the extent of damage reaches the value for residual strength evaluation) provides a practical basis for development of the inspection programs and procedures described in paragraph AC 29 MG 11g(8). For multiengine load paths, a minimum of three inspection intervals is recommended between the initially detectable damage time and the time when residual strength is reduced to design limit load by crack growth. For single element structures, a minimum of four inspection intervals is recommended. The repeated loads should be defined in the loading, temperature, and humidity spectra. The loading conditions should take into account the effects of structural flexibility and rate of loading where it is significant. (v) For flaw tolerance to achieve an improvement in safety over safe-life for composite structures, the procedures of Paragraph 7a of AC 20-107A are recommended. For flaw tolerance to achieve and improvement in safety over simple safe-life for metallic structures, the following testing criteria are recommended to supplement flaw tolerant safe-life analysis using flawed specimen S-N curves or crack growth analyses using appropriate stress intensity factors and da/dn data from coupon tests: (A) Flaw Tolerant (Enhanced) Safe-Life Testing Criteria. Test full-scale specimens with flaws or a mix of full-scale and coupon specimens with flaws to obtain S-N data. Plot the data as shown in figure AC 29 MG 11-1. Utilize the S-N data andAC 29-2C 9/30/99

Page MG 11 - 16 loading spectrum of paragraph AC 29 MG 11 f(2) in substantiating a crack-free life or in arriving at a replacement time by cumulative damage analysis means. Where practical (as allowed by the number of damaging cycles in the loading spectrum), spectrum testing may be used for fatigue substantiation in lieu of S-N testing followed by analysis. The replacement time established should be included in the airworthiness limitations section of the document established under § 29.1529. (B) Fail-Safe Crack Growth (Fracture Mechanics) Substantiation. Test two or more specimens to obtain crack propagation data using either a realistic load spectrum or an accelerated load (spectrum or single) associated with the use of propagation theory and data after cracks have been initiated. Unless a more rational method with an equivalent level-of-safety is applied for, the following methods of setting inspection intervals should be applied. In all cases, the inspection methods and intervals should adequately consider variables such as inspectability, type of inspection, crack growth behavior, and other scheduled maintenance considerations. (1) For single element (load path) structure, plot the data and set the inspection as shown in figure AC 29 MG 11-2. (i) Set the initial inspection at L 1 /3. (ii) Set the repetitive inspection intervals at L 2 /4.9/30/99 AC 29-2C

Page MG 11 - 17AC 29-2C 9/30/99

Page MG 11 - 18 (2) For multi-element load path structure: (i) Test all paths simultaneously with a 0.05-inch thick crack (or size detectable by manufacturing quality control procedures) in the critical element at the start of tests. (ii) Note when field detectable cracking occurs. (iii) Note when one element fails. (iv) Note when the residual strength of the remaining elements decreases to limit load due to crack growth. (v) From figure AC 29 MG 11-3, set initial inspection at L 1 /3. (vi) Set repetitive inspection intervals at (L 2+ L R )/3. NOTE: If partial failure of the critical element is not detectable, then L 2becomes zero.9/30/99 AC 29-2C

Page MG 11 - 19AC 29-2C 9/30/99

Page MG 11 - 20 (C) Other test and inspection programs may be used rather than those of paragraph AC 29 MG 11 g(7)(v)(A) and paragraph AC 29 MG 11 g(7)(v)(B), if shown to have comparable or better probability of assuring that a catastrophic fatigue failure will not occur. (vi) The flaw tolerance characteristics can also be shown analytically by reliable or conservative methods such as the following: (A) By demonstrating quantitative relationships with structure already verified as damage tolerant; (B) By demonstrating that the damage would be detected before it reaches the value for residual strength evaluation; or (C) By demonstrating that the repeated loads and limit load stresses do not exceed those of previously verified designs of similar configuration, materials, and inspectability. (8) Inspection. Detection of flaws before they become dangerous is the ultimate control in ensuring the flaw tolerance characteristics of the structure. Therefore, the applicant should provide sufficient guidance information to assist operators in establishing the frequency, extent, and methods of inspection of the critical structure, and this kind of information should, under § 29.571(a)(2), be included in the maintenance manual required by § 29.1529. Due to the inherent, complex interactions of the many parameters affecting flaw tolerance, such as operating practices, environmental effects, load sequence on flaw growth, and variations in inspection methods, related operational experience should be taken into account in establishing inspection procedures. Comparative analysis can be used to guide the changes from successful past practice, when necessary. Therefore, maintenance and inspection requirements should recognize the dependence on experience and should be specified in a document that provides for revision as a result of operational experience, such as the one containing the operator’s FAA/AUTHORITY-approved structural inspection program developed through the Maintenance Review Board (MRB) procedures for FAR Part 121 operators.

h. COMBINATION OF REPLACEMENT TIME AND FLAW GROWTH EVALUATION. It may be possible to extend the replacement time of safe-life components that exhibit limited flaw tolerance capability by using a combination of the safe-life and flaw growth characteristics as described elsewhere in paragraph AC 29 MG 11 and by assigning both a replacement time and inspection period to these components. The replacement time may then be based on the combined probability of not initiating a fatigue crack at or before the replacement time and the probability that the crack, if initiated, will be detected prior to catastrophic failure or loss of limit load (or maximum attainable load, whichever is less) carrying capability. The probability of detection should be based on consideration of the inspection effectiveness, the inspection intervals, and the fatigue life remaining after an obvious9/30/99 AC 29-2C

Page MG 11 - 21 partial failure. A lower strength reduction factor commensurate with this probability of detection may then be used in the determination of the replacement time.AC 29-2C 9/30/99

Page MG 11 - 22 Attachment 1 to AC 29 MG 11 FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED

1. GROUND CONDITIONS. a. Normal start. b. Rapid increases of RPM on ground to maximum power-on RPM of main rotor. c. Taxiing with full cyclic control. d. Landing run (if applicable). e. Braking (if applicable). f. Normal shutdown. g. Special ground checks (if applicable). 2. IN-GROUND-EFFECT (IGE) MANEUVERS. a. Hovering. (1) Steady with rotor at maximum side of RPM tolerance. (2) Steady with rotor at minimum side of RPM tolerance. (3) 90° right turn. (4) 90° left turn. (5) Control reversal. (i) Longitudinal. (ii) Lateral. (iii) Rudder. (6) Sideward flight. (i) Right. (ii) Left. (iii) Rearward flight. b. Maneuvering. (1) Jump takeoff. (2) Normal takeoff and accelerate to climb airspeed. (3) Normal approach and landing. (i) Multiengine. (ii) One-engine-inoperative. (4) Full autorotational landing.9/30/99 AC 29-2C

Page MG 11 - 23 Attachment 1 to AC 29 MG 11 FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED (continued)

3. FORWARD FLIGHT-POWER ON. a. Level flight. (1) 40 percent V H .

(i) Minimum side of main rotor RPM tolerance (RPM +) (ii) Maximum side of main rotor RPM tolerance (RPM -) (2) 60 percent V H .

(i) (RPM +) (ii) (RPM -) (3) 80 percent V H .

(i) (RPM +) (ii) (RPM -) (4) V H .

(i) (RPM +) (ii) (RPM -) (5) V NE .

(i) (RPM +) (ii) (RPM -)

b. Maneuvers. (1) Full power climbs.

(i) All engines operative. (ii) One-engine-inoperative. (2) Cyclic pull-ups.

(i) 60 percent V H . (ii) 90 percent V H . (3) Normal acceleration from climb airspeed to 90 percent V H . (4) Turns.

(i) Right at 60 percent V Hand 90 percent V H . (ii) Left at 60 percent V Hand 90 percent V H . (5) Control reversals at 90 percent V H .

(i) Longitudinal. (ii) Lateral. (iii) Rudder. (6) Deceleration from 90 percent V Hto descent airspeed. (7) Part power descent.

(i) All engines. (ii) One engine out.AC 29-2C 9/30/99

Page MG 11 - 24 Attachment 1 to AC 29 MG 11 FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED (continued)

4. POWER TRANSITIONS. a. All engines operating to one engine out. (1) In full power climb. (2) At 90 percent V H . b. One engine out to all engines operating in powered descent. c. All engines operating to autorotation. (1) At 60 percent V H . (2) At maximum forward transition speed. d. Stabilized autorotation to all engines operating at normal autorotation airspeed.

5. AUTOROTATION.

a. Stabilized. (1) At 70 percent V NE . (2) At NE . b. Turns at 70 percent and 100 percent V NE . (1) Right. (2) Left. c. Cyclic pull-up.

d. Control reversals. (1) Longitudinal. (2) Lateral. (3) Rudder. The flight and ground conditions of this table may be used as practical in developing operating load spectra for the range of projected aircraft usage such as: (1) Short missions with high-power cycles and ground-air-ground (GAG) cycles (such as used in logging operations). (2) Medium range missions of 1 to 2 hours that include three or more GAG or LO-HI power cycles per mission.9/30/99 AC 29-2C

Page MG 11 - 25 Attachment 1 to AC 29 MG 11 FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED (continued) (3) Long range missions that consider the longest practical range of the aircraft and consider two or more GAG (or high power) cycles per mission (such as used for remote offshore drilling platforms). If aircraft usages that produce a large number of high-power cycles (approximately 50 per hour) or high GAG cycles (approximately 25 per hour) such as logging operations are not included in the fatigue tolerance substantiation, this should be noted in the RFM and Maintenance Manual airworthiness limitations section (reference paragraph AC 29 MG 11 f(1)).