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
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 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. This approach will minimize the risk associated with substantiation of the full-scale article. It is strongly recommended that a certification team approach, involving fabrication, quality, and engineering specialists from both the applicant and FAA/AUTHORITY, be used for composite structural substantiation.
The team should assure that permanent documentation of the building block approach in the form of reports or other FAA/AUTHORITY-acceptable documents are included in the certification data package. The documentation includes but is not limited to the structural substantiation reports (both analysis and test), manufacturing processes and quality control, and Instructions for Continued Airworthiness (maintenance, overhaul, and repair manuals). FAA/AUTHORITY engineering approves the Airworthiness Limitations Section of the maintenance manual. Engineering practices for many of the areas identified below are available in MIL-HDBK-17.
(1) The first area is the fabrication requirements of § 29.605:
(i) The quality system should be developed considering the critical engineering, manufacturing, and quality requirements along with 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. Some important concepts of AC 21-26 are included below.
(ii) Specific allowable defect limits on, for example, fiber waviness, warp defects, fill defects, porosity, hole edge effects, edge defects, resin content, large area disbonds , 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 variability of the manufacturing process to determine the worst-case effects (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) Page MG 8 - 9
AC 29-2C, Chg 2 4/25/06 of this AC paragraph) 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-17.
These methods can also be used, periodically for qualification status renewals of existing material systems and suppliers.
(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 systems 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 (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 under their approved production inspection, fabrication inspection, and quality systems .
(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 under their approved production inspection, fabrication inspection, and quality systems.
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(D) Statistical Validation Level. It is necessary to maintain the minimum required statistical validation level of the quality 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 must 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 composite 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” and “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.
Although MIL-HDBK-17 does have some "B" basis allowables available in Volume 2, company testing is required for "A" basis and other "B" basis material systems not listed. 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 that 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 Page MG 8 - 11
AC 29-2C, Chg 2 4/25/06 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 methodolog y of MIL-HDBK-17 should be implemented, i.e., the wet glass transition temperature (Tg) should be 50° F higher than the maximum structural temperature (see definition in paragraph e(24) of this AC paragraph). 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 or experience or both should be established and approved early in the certification process.
(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 such that specified impact damage considerations are defined and properly limited. The effects of the strain levels may be established for each composite material using small-scale characterization tests and then 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 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 Page MG 8 - 12