CHAPTER 3
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 7. STRUCTURAL CONDITION INDICATORS.
a. Related Sections. § 29.301 - Loads; § 29.305 - Strength and Deformation;
§ 29.571 - Fatigue Evaluation of Flight Structure; § 29.1301 - Function and Installation;
§ 29.1309 - Equipment, Systems and Installations; § 29.1321 - Arrangement and Visibility; § 29.1322 - Warning, Caution, and Advisory Lights; § 29.1355 - Distribution System; § 29.1503 - Airspeed Limitations: General; and § 29.1529 - Instructions for Continued Airworthiness.
b. Background.
(1) Structural condition indicators have been used on rotorcraft for several years in two main programs: as part of the basic type design and as part of airworthiness directive (AD) action. When approved as part of the basic type design, only limited “credit” has been given for the installation of structural condition indicators;
i.e., components provided with a structural condition indication system were required to be designed to § 29.571 “safe-life” criteria considering the structural condition indicator system inoperative. So- called “nonhazard” approvals were granted. When used as part of the mandatory actions of ADs, structural condition indicators have had a degree of “credit” recognized, primarily in the recognition of “fail-safety” provided by the indicator system.
(2) Since structural condition indicators have been used during both original type design and AD issuance, and since there is movement toward increased damage tolerance in rotorcraft design, policy concerning condition indicator use is considered appropriate.
c. At present, the use of structural condition indicators alone on new type designs is not considered an acceptable substitute for providing the necessary safe life for each component. However, areas which may be considered when approving these indicators for fail-safety credit are delineated in the following paragraphs.
d. What, how, when, where, and who of structural condition indicators.
(1) Indication of what?
(i) Previous structural condition indicators have primarily been used for crack detection. Several types of through- the-thickness crack detection systems are currently in use. Two types which detect changes in pressure in an instrumented chamber due to gas movement through a cracked wall are known as the blade Page MG 7 - 1
inspection method (BIM) system and the integral spar inspection system (ISIS). These systems can only detect full-depth cracks which are large enough to allow loss (of gain) of pressure from the instrumented chamber. This presents a limitation since full -depth cracks may be fast growing before detection. Another through-t he-thickness crack method is a pressurized, dyed fluid or oil system to detect through cracks in specially designed bolts (NASA patent), spindles, pins, or other closed chamber mechanical equipment.
(ii) Surface cracks can be found by systems such as surface-mounted crack detection wires. These systems would allow a greater safe crack growth period for assuring safe landing after detection than the through-crack-detection systems, but they have been used little in operations because of significant limitations; e.g., complexity of installation, durability problems, limited areas of coverage, and strain level limitations.
(iii) Some aircraft have had mast moment indicators or other load indicators to help prevent the pilot from inadvertently applying a high load to the instrumented system or to help the pilot reduce the load by control movements. These load indicators only indirectly give indications of structural condition; therefore, only limited “credit” is allowed for this use. “Credit” is limite d in that the fatigue life substantiations of § 29.571 should consider a reasonable number of excursions into the higher ranges established for the load indicator, and special inspections, rework, or replacement instructions should be provided for any strength degradation associated with high range excursions.
(2) How indicated?
(i) Current BIM systems use two types of indicators. The visual blade inspection method (VBIM) uses a gauge mounted on the blade which must be read visually by maintenance personnel while the aircraft is parked. The cockpit blade inspection method (CBIM) uses lights mounted in the cockpit which may be monitored by the crew. Other pressurized chambers have used dyes or oils to improve visual inspection effectiveness. Mast moment indicators and other load indicators use instruments with marked ranges and needles.
(ii) No specific types of load indicators are required by the FAA/AUTHORITY but the type used should be evaluated for accuracy, readability, and overall effectiveness. Paragraphs AC 29 MG 7(e) and (f) cover, in more detail, the use of structural condition indicators.
(3) When indicated? Structural condition indicators are used before flight, during flight, and for normal maintenance inspections. Paragraphs AC 29 MG 7 (e) and
(f) contain guidance for cockpit-mounted instruments which are monitored during flight.
Indicators used for normal maintenance inspections are the preferred type since they can be scheduled to allow the most effective use of available maintenance personnel of well-equipped maintenance facilities and of parts available.
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(4) Where indicated? Indications on the component are provided by VBIM systems and by systems utilizing dye or colored oil leakage. Cockpit -mounted lights and gauges may be used for certain critical structures which require frequent, but simple, checks. Maintenance panel locations (cabin, equipment bay, etc.) are the preferred locations for use in routine maintenance.
(5) Who reads indicators? The flightcrew, of necessity, monitors indicators mounted in the cockpit for use during flight. Gauges with ranges of values representing mast bending moments or other structural loads are monitored by the flightcrew, as necessary, to reduce or to prevent control operations from imposing excessive loads or to prevent too many high load applications. Maintenance personnel are generally responsible for reading component-mounted indicators and for monitoring indicators which are mounted on maintenance panels. The before- flight checks may be conducted by maintenance personnel or by flightcrew in certain cases (i.e., cockpit-mounted gauges or “push-to-test” checks).
e. Actions required by indicators .
(1) On-ground indications. Indications noted on the ground should be followed by a functional check of the indication system as provided for by its design. If indications persist after the system has been checked and found to be functional, further inspection of the affected component(s) should be conducted for damage assessment.
Any damage found as a result of the detailed inspections should be repaired or replaced as appropriate.
(2) In-flight indications.
(i) Indications used for in-flight monitoring have in the past been used for two main reasons: to provide a structural load display (such as mast bending moment) and to help resolve a service problem (CBIM systems have been used to supplement conventional inspection methods in blind areas).
(ii) Structural load display systems should not be used instead of correcting deficient designs. Structural load display systems are appropriate for use in locating control positions, such as the cyclic stick, under transient conditions such as slope landings and hover in sidewinds, but structural load display systems are not considered appropriate for routine operations such as climbout or cruise with constant attention required by the flightcrew. If the load indicator provides a needed tool to the pilot in limited types of operations and does not significantly add to pilot workload otherwise, its use can be considered.
(iii) In the past, certain service problems have been solved by adding in-flight indicators such as CBIM systems. When retrofit of the affected structure is impossible or impractical, and when conventional inspection techniques are shown to be inadequate by themselves, CBIM or similar systems may be the only practical Page MG 7 - 3
solution, despite the increase in pilot workload and the potential for problems caused by overreaction by the pilot to a structural fault indication. When used for correction of service difficulties, the structural condition indicator system should be accompanied by clear, concise crew directions to prevent possible catastrophic overreaction. Load reduction measures such as rotor speed changes, airspeed reductions, altitude changes, etc., should be clearly provided, if needed. Crack propagation time from indication should be sufficient to allow continued safe flight to a safe landing area. For new designs, CBIM or similar systems which add to the pilot’s workload are considered inappropriate. Proper redesign to provide the needed safe life, fail -safety, and inspectability is considered the appropriate action.
f. Complementary considerations of structural condition indicator use.
(1) Two basic programs are commonly used for approval of structural condition indicators. Basic type certification procedures are used for mast moment indicators and similar systems, and AD’s (with appropriate type design changes) are used for CBIM systems which require pilot attention and corrective action when an indication of a structural fault is detected.
(2) The fatigue substantiation required by §§ 27.571 and 29.571 should consider a conservative number of excursions into the high load range monitored by a structural condition indicator such as a mast moment indicator. Static strength should not be adversely affected by a single excursion into the high load range monitored by the indicator.
(3) Complementary design provisions should accompany the use of a structural condition indicator system. Redundancy of load paths and inspection systems and indicator system failure analyses should be provided, as necessary, to meet the requirements of § 29.1309. The life remaining after the indicator system detec ts a structural failure should be calculated (with test verification), and compatible inspection and/or overhaul programs should be provided.
(4) The FAA/AUTHORITY approval of a structural condition indicator system requires evaluation by the airframe, systems and equipment, and flight test specialists.
The airframe specialist has the responsibility to review effects of structural condition indicator system use on aircraft loads, strength and deformation, and structural fatigue evaluation as well as the instructions for continued airworthiness. The systems and equipment specialist needs to evaluate the system for function and installation as well as the reliability requirements of § 29.1309. Flight test evaluation of the instruments’ arrangement and visibility, effect on crew workload, and possible changes for RFM is also needed. Care should be exercised to assure that responsibilities are not given to the flightcrew which would be more appropriately handled by a redesign or by the maintenance personnel. Early coordination between all specialists is necessary to prevent delays from last minute design changes.
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CHAPTER 3
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 8. (Amendment 29-42) SUBSTANTIATION OF COMPOSITE ROTORCRAFT STRUCTURE.
a. Reference. 14 CFR §§ 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 procedur es 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 FAA/AUTHORITY.
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 st rength substantiation for both metal and composite structure is essentially identical to that for fixed wing structure once basic loads have been determined. However, rotorcraft st ructural fatigue substantiation for composites is significantly different from fixed wing fatigue substantiation. Since AC 20-107A, as dev eloped, 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 component fatigue tests for rotorcraft programs rather than just one full-scale test. Also, constant amplitude, Page MG 8 - 1
AC 29-2C, Chg 2 4/25/06 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 AC 29-2C, Chg 1, MG 11, “Fatigue Tolerance Evaluation of Transport Category Rotorcraft Metallic 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 AC 27-1B, Chg 1, 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., MIL-HDBK-5 for metals, MIL-HDBK-17 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 natur e, 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. Current 14 CFR Page MG 8 - 2
Parts 27 and 29 contain many sections (reference paragraph a above) 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 14 CFR 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.
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AC 29-2C, Chg 2 4/25/06
(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.
(11) FIBER. A s ingle 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 (e.g., 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.
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(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., each 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 material, 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 that will involve complete curing. Pre-preg is usually drapabl e, 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 jo ining 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 lengths 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 micro inches/per inch (e.g., .003 in/in equals 3000 micro inches/inch).
(32) SYMMETRICAL LAMINATE. A composite laminate in which the ply orientation is symmetrical about the laminate midplane.
(33) TAPE. Hot melt impregnated fi bers 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.
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AC 29-2C, Chg 2 4/25/06
(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 lengt h of the fabric (in the 0° direction) and being crossed by the fill yarns (90° fibers).
(37) WORK LIFE. The period duri ng which a compound, after mixing with a catalyst, solvent, or other compounding constituents, remains suitable for its intended use.
(38) CATASTROPHIC FAILURE. Any stru ctural failure, which results in death, severe injury, or loss of the aircraft.
(39) FATIGUE TOLERANCE. The capability of struct ure to continue functioning without catastrophic failure after being subjected to fatigue (repeated) loads expected during operation of the rotorcraft. Fatigue tolerance should be achieved by flaw tolerance design, or if impractical, safe-life design, or a combination.
(40) SAFE-LIFE. The capability of as-manufactured structure as shown by tests, or analysis based on tests, not to initiate fatigue cracks during the service life of the rotorcraft or before an established replacement time.
(41) FLAW TOLERANCE. The capability of rotorcraft structure to achieve fatigue tolerance accounting for the presence of flaws and damage that may occur in manufacturing and service use. Flaw tolerance can be achieved by either flaw tolerance safe-life or fail-safe designs. The term “Damage Tolerance” is frequently used to describe the ability of a structure to tolerate the effects of flaws and damage;
however, the terminology of § 29.571, Amendment 28, is used in this AC to maintain consistency.
(42) FLAW TOLERANT SAFE-LIFE. The capability of as-manufactured structure, with expected flaws, as shown by tests or analysis based on tests, not to initiate fatigue cracks or flaw/damage growth during the service life of the rotorcraft or before an established replacement time.
(43) FAIL-SAFE. The capability of structure remaining after a partial failure to withstand design limit loads without catastrophic failure within an inspection period.
(44) MULTIPLE LOAD PATH. Struct ure providing two or more separate and distinct paths of structure that will carry limit load after complete failure of one of the members.
(45) ACTIVE MULTIPLE LOAD PATH. Structure providing two or more load paths that are all loaded during operation to a similar load spectrum.
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(46) 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.
(47) ACCIDENTAL DAMAGE FLAWS. Discrete damage that may occur in service use or in manufacturing due to impacts or collisions, such as dents, scratches, gouges, abrasions, disbonds, splintering, and delaminations.
(48) MANUFACTURING-RELATED FLAWS. Intrinsic imperfections related to manufacturing operations, processing, or assembly such as voids, gaps, porosity, inclusions, fiber dislocation, disbonds, and delaminations.
(49) FATIGUE/ENVIRONMENTAL FLAWS. Structural damage related to fatigue or environmental effects such as delaminations, disbonds, splintering, or cracking.
(50) DESIGN LIMIT LOADS. The ma ximum loads to be expected in service, as defined by § 29.301(a).
(51) 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.
(52) RESIDUAL STRENGTH. The strength retained for some period of unrepaired use after a failure or partial failure due to fatigue or accidental or discrete source of damage.
(53) PRINCIPAL STRUCTURAL ELEMENT (PSE). A structural element that contributes significantly to the carrying of flight or ground loads and whose failure can lead to catastrophic failure of the rotorcraft.
(54) 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).
(55) POINT DESIGN. An element or detail of a specific design that is not considered generically applicable to other structure for the purpose of substantiation
(e.g., lugs and major joints). Such a design element or detail can be qualified by test or by a combination of test and analysis.
(56) ELEMENT. A generic element of a more complex structural member
(e.g., skin, stringers, shear panels, sandwich panels, joints, or splices).
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AC 29-2C, Chg 2 4/25/06
(57) DETAIL. A non-generic structural element of a more complex structural member (e.g., specific design configurated joints, splices, stringers, stringer runouts, or major access holes).
(58) SUBCOMPONENT. A major thre e-dimensional structure, which can provide complete structural representation of a section of the full structure (e.g., stub box, section of a spar, wing panel, wing rib, body panel, or frames).
(59) COMPONENT. A major section of the airframe structure (e.g., wing, body, fin, horizontal stabilizer), which can be tested as a complete unit to qualify the structure.
(60) ENVIRONMENT. External , nonaccidental 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).
f. Related Regulatory and Guidance Material.
Document Title FAA Order 8110.9 Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopter and Other Power Transmission Systems AC 27-1B, Chg 1, Fatigue Evaluation of Rotorcraft Structure MG 11 AC 20-107A Composite Aircraft Structure AC 21-26 Quality Control for the Manufacture of Composite Materials MIL-HDBK-17 Composite Material Handbooks AC 29-2C, Chg 1, Fatigue Tolerance Evaluation of Transport Category MG 11 Rotorcraft Metallic Structure (Including Flaw Tolerance) DOT/FAA/CT-86/39 Whitehead, R.S., Kan, H.P., Cordero, R., and Seather, R., “Certification Testing Methodology for Composite Structures”, October 1986.
g. PROCEDURES FOR SUBSTANTIATION OF ROTORCRAFT COMPOSITE STRUCTURE. The composite structures ev aluation 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) fatigue tolerance evaluation (including tolerance to flaws); (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, co-cured assemblies, bonded assemblies, etc.) should Page MG 8 - 8