AC 29-2C, Chg 4 12/2/2011 be contained within the ALS of the ICAs unless they are structural inspection intervals or related structural inspection procedures approved under § 29.571.
(10) Approved Equivalent Means. The requirement includes the possibility that in place of setting retirement times or inspections for damage, some other means may be used. All proposals for ‘equivalent means’ must be submitted to the FAA for approval. Potentially equivalent means to inspection include, but are not limited to:
(i) Indirect detection of damage used to establish a period of safe operation for a structure with the damage present. In this case, the detection is based on the effect of the damage, which may be recognized through:
(A) A warning in flight or during maintenance from a specific feature, sensor, or health monitor, including: oil analysis, chip detector, crack detection wire or foil, health monitoring, fluid leaks or pressure change in a sealed chamber; or by (B) Pilot sensitivity to a change in the rotorcraft’s behavior (such as poor blade tracking, noise generation, vibration generation) provided it is well defined and does not require exceptional piloting skills to recognize these behaviors.
(ii) In all cases, an adequate level of residual strength is demonstrated for the period of operation concerned. Generally, limit load will be considered the minimum residual strength requirement. However, load levels less than the critical limit load conditions may be acceptable for consideration of obvious damage sustained in flight and for the completion of that flight only, provided it allows for continued safe flight and landing.
(iii) Two instances are considered here where it may not be necessary to provide a retirement time in the ALS of the ICAs. However, this does not preclude the investigation of fatigue behavior throughout the life of the rotorcraft or of the part if longer.
(A) When fatigue cracking occurs, or is expected to occur, for a specific PSE while in service, then the first approach allows the PSE to operate until the damage is found. Therefore, the inspection must find the damage prior to loss of ultimate load capability. This approach may not be appropriate for a single load path structure. For such a process to be safe, the behavior of the part and associated parts that influence its fatigue behavior must be substantiated for as long as they remain in service. All potential failure modes throughout the life of the rotorcraft must be identified and shown to be consistent, repeatable and addressed by the inspection program. In order to meet the intent of the new fatigue tolerance requirements, a high probability of ultimate load capability is required throughout the lifetime of the component. Therefore, for cracks or other damage that are allowed or highly likely to exist, ultimate load capability should be substantiated for that damage and any growth that may occur during the subsequent inspection period.
Page C - 130
(B) It may be acceptable that a PSE does not have a specific retirement time when the fatigue tolerance of the part, including any damage not controlled by an acceptable inspection program, has been demonstrated to be in excess of the rotorcraft design life to such an extent that no safety benefit arises from imposing that requirement.
(11) Supplemental Procedures.
(i) The requirement states that if inspections, for any of the damage types identified during the threat assessment, cannot be established within the limitations of geometry, inspectability or good design practice, then supplemental procedures must be established that will minimize the risk of each of these types of damage being present or leading to catastrophic failure. When assessing good design practice, measures such as improved protection against impact, scratches, and corrosion should already have been considered. If the part cannot be redesigned to reduce the acquisition and influence of damage, then supplemental procedures should be introduced.
(ii) Supplemental procedures that should be considered include, but are not limited to:
(A) Specifying shorter than usual calendar inspection intervals to reduce the probability of occurrence and the extent of the damage.
(B) Improving control of maintenance processes associated with the component and damage type, such as by providing specifically designed tooling and requiring additional quality checks after each operation is performed.
(C) Introducing an overhaul program.
(D) Restricting the allowable repair limits for the part.
(E) Modifying the PSE design based on service experience if this shows the original design assumptions to be overly conservative with respect to demonstrating impracticality at certification.
(F) Specifying a conservative inspection interval, if the calculated interval cannot be established and there are no other alternatives.
Page C - 131
AC 29-2C, Chg 4 12/1/2011 AC 29.573. § 29.573 (Amendment 29- 54) DAMAGE TOLERANCE AND
FATIGUE EVALUATION OF COMPOSITE ROTORCRAFT
STRUCTURES
a. Purpose. This advisory material provides an acceptable means of compliance with the provisions of § 29.573, Amendment 29-54, Title 14 of the Code of Federal Regulations (CFR) dealing with the damage tolerance and fatigue evaluation of transport category composite rotorcraft structures. Paragraph f.(6) specifically addresses the advisory guidance applying to damage tolerance and fatigue evaluation as required by § 29.573, Amendment 29-54. Some information contained in AC 29- 2C, MG 8 (Amendment 29-42) is repeated and updated, as appropriate, to preserve the “building block” approach for analyses of composite rotorcraft structure for compliance to § 29.573, Amendment 29-54 . (Supplemental guidance can be found in AC 20- 107, “Composite Aircraft Structure.”) These procedures address the substantiation requirements for composite material system constituents, composite material systems, and composite structures common to rotorcraft. A uniform approach to composite structural substantiation is desirable, but it is recognized that in a continually developing technical area, which has diverse industrial roots both in aerospace and in other industries, variations and deviations from the procedures described here may be necessary. Deviations from this advisory material should be coordinated in advance with the Rotorcraft Directorate.
b. Special Considerations. Since rotorcraft structure is configured uniquely and is inherently subjected to severe cyclic stresses, special consideration is required for the substantiation of all rotorcraft structure, including composites. This special consideration is necessary to ensure that the level of safety intended by the current regulations are attained during the type certification process for all structure with special emphasis on composite structure because of its unique structural characteristics, manufacturing quality and operational considerations, and failure mechanisms.
c. Background.
(1) Historically, rotorcraft have required unique, conservative structural substantiation because of unique configuration effects, unique loading considerations, severe fatigue spectrum effects, and the specialized comprehensive fatigue testi ng required by these effects. Rotorcraft structural static strength substantiation for both metal and composite structure is essentially identical to that for fixed wing structure once basic loads have been determined. However, rotorcraft structural fatigue substantiation is significantly different from fixed wing fatigue substantiation. Since AC 20-107, as developed, applies to both fixed wing aircraft and rotorcraft, it, of necessity, was finalized in a broad generic form. Accordingly, a need to supplement AC 20-107 for rotorcraft was recognized during type certification programs. One significant difference in traditional rotorcraft fatigue substantiation programs and fixed wing fatigue programs is the use of multiple component fatigue tests for rotor craft programs rather than just one full-scale test. Also, constant amplitude, accelerated load tests are typically used rather than spectrum tests because of the high frequency Page C - 132
loads common to rotorcraft operations. These rotorcraft fatigue tests have traditionally involved the generation of stress versus life or cycle (S-N) curves for each critical part (most of which are subjected to the cyclic loading of the main or tail rotor system) using a monotonic (sinusoidal) fatigue spectrum based on maximum and minimum service stress values. Unless configuration differences or flight usage data dictate otherwise, the monotonic fatigue spectrum’s period is typically based on six ground-air-ground (GAG) cycles for each flight hour of operation. The S -N curves for the substantiation of each detailed part are typically generated by plotting a curved line through three data points (see AC 29-2C, AC 29 MG 11, “Fatigue Tolerance Evaluation of Transport Category Rotorcraft Metallic Structure”). The three data points selected are a short specimen life (low-cycle fatigue), an intermediate specimen life and a long specimen life (high-cycle fatigue). Each raw data point is generated by monotonically fatigue testing at least two full-scale parts to failure or run out for each data point on the S -N curve.
The raw data point values are then reduced by an acceptable statistical method to a single value for plotting to ensure proper reliability of the associated S-N curve.
Order 8110.9, “Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopter and Other Power Transmission Systems” and AC 27-1B, AC 27 MG 11, “Fatigue Evaluation of Rotorcraft Structure”, contain comprehensive discussions of the S-N curve generation process. The rotorcraft S -N curve process contrasts sharply with the fixed wing process of using a single full-scale fatigue article (usually an entire wing or airframe, which constitutes a single full-scale assembly data point), generic material or full-scale assembly S-N data (e.g., Metallic Materials Properties Development and Standardization (MMPDS), formerly the MIL-HDBK-5 for metals; Composites Materials Handbook-17 (CMH-17), formerly the MIL-HDBK-17 for composites; or AC 23-13, “Fatigue, Fail-Safe, and Damage Tolerance Evaluation of Metallic Structure for Normal, Utility, Acrobatic, and Commuter Airplanes”, which replaced AFS-120-73-2 for full-scale assemblies), a non-monotonic spectrum, and relatively large scatter factors to verify or determine the design fatigue life of the full-scale airplane.
(2) Additionally, rotorcraft have employed and mass-produced composite designs in primary structure (typically main and tail rotor blades) since the early 1950’s.
This was 10 or more years before composites were type certificated for primary fixed-wing structure in either military or civil aircraft applications (with some notable limited production exceptions, such as the Windecker fixed wing aircraft). In any case, the early 1950 period was well before a clear, detailed understanding of composite structural behavior (especially in the areas of macroscopic and microscopic failure mechanisms and modes) was relatively common and readily available in a usable format for the average engineer working in this field. It also predated the initial issuance of AC 20-107. Currently, much composite design information is proprietary, either to government, industry or both, and many data gathering methods have not been completely standardized. Consequently, a significant variation from laboratory to laboratory in material property value determination methods and results can exist. The early rotor blade designs (as well as current designs) are by nature relatively low strain, tension structure designs. Also, by nature, these designs are not damage or flaw critical. Thus, by circumstance as much as design, early composite rotor blade and other composite rotorcraft designs incorporated an acceptable fatigue tolerance level of Page C - 133
AC 29-2C, Chg 4 12/1/2011 safety. In the 1980’s, more test data, analytical knowledge, and analytical methodology became available to more completely substantiate a composite design. Current 14 CFR parts 27 and 29 contain many sections to be considered in substantiating composite rotorcraft structure. This advisory material provides the current or updated information from AC 29-2C, MG 8, Amendment 29-42 to supplement the general guidance of AC 20-107 and provides compliance guidance for the requirements of § 29.573 Amendment 29-54 for rotorcraft composite structure.
d. Definitions. The following basic definitions are provided as a convenient reading reference. CMH -17, and other sources, contain more complete glossaries of definitions.
(1) A-Basis Allowable. The “A” mechanical property value is the value above which at least 99 percent of the population of values is expected to fall, with a confidence level of 95 percent.
(2) Accidental Damage. Discrete damage, which may occur in service use or in manufacturing due to impacts or collisions, such as dents, scratches, gouges, abrasions, disbonds, splintering, and delaminations.
(3) Active Multiple Load Path. Structure providing two or more load paths that are all loaded during operation to a similar load spectrum.
(4) Allowables. Both A -basis and B-basis values statistically derived and used for a particular composite design.
(5) As-Manufactured. Product or component that has passed the applicable quality control process and has been found to conform to the approved design within the allowable tolerances.
(6) Autoclave. A closed apparatus usually equipped with variable conditions of vacuum, pressure, and temperature. It is used for bonding, compressing or curing materials.
(7) B-Basis Allowable. The “B” mechanical property value is the value above which at least 90 percent of the population of values is expected to fall, with a confidence level of 95 percent.
(8) Balanced Laminate. A composite laminate in which all laminae at angles other than 0° occur only in ± pairs (not necessarily adjacent).
(9) Bond. The adhesion of one surface to another, with or without the use of an adhesive as a bonding agent.
(10) Catastrophic Failure. An event that could prevent continued safe flight and landing.
Page C - 134
(11) Cocure. The process of curing several different materials in a single step.
Examples include the curing of various compatible resin system pre-pregs, using the same cure cycle, to produce hybrid composite structure or the curing of compatible composite materials and structural adhesives, using the same cure cycle, to produce sandwich structure or skins with integrally molded fittings.
(12) Component. A major section of the airframe structure (e.g., wing, fin, body, horizontal stabilizer), which can be tested as a complete unit to qualify the structure.
(13) Coupon. A small test specimen (e.g., usually a flat laminate) for evaluation of basic lamina or laminate properties or properties of generic structural features (e.g., bonded or mechanically fastened joints).
(14) Cure. To change the properties of a thermosetting resin irreversibly by chemical reaction (i.e., condensation, ring closure, or addition). Cure may be accomplished by addition of curing (crosslinking) agents, with or without a catalyst, and with or without heat.
(15) Damage. A generic term for structural anomalies caused by manufacturing (processing, fabrication, assembly or handling) or service usage. Trimming, fastener installation, or foreign object impact are potential sources of damage, along with fatigue and environmental effects.
(16) Damage Tolerance. The attribute of the structure that permits it to retain its required residual strength for a period of use after the structure has sustained a given level of fatigue, corrosion, accidental or discrete source damage.
(17) Damage Tolerant Fail-Safe. The capability of structure remaining after a partial failure to withstand design limit loads without catastrophic failure within an inspection period.
(18) Damage Tolerant Safe Life. Capability of structure with damage present to survive expected repeated loads of variable magnitude without detectable damage growth and to maintain ultimate load capability throughout service life of the rotorcraft.
(19) Delamination. The separation of the layers of material in a laminate.
(20) Design Limit Loads. The maximum loads to be expected in service, as defined by § 29.301(a).
(21) Detail. A non- generic structural element of a more complex structural member (e.g., specific design configured joints, splices, stringers, stringer runouts, or major access holes).
Page C - 135
AC 29-2C, Chg 4 12/1/2011
(22) Disbond. A lack of proper adhesion in a bonded joint. This may be isolated or may cover a majority of the bond area. It may occur at any time in the cure or subsequent life of the bond area and may arise from a wide variety of causes.
(23) Element. A generic part of a more complex structural member (e.g., skin, stringers, shear panels, sandwich panels, joints, or splices).
(24) Environment. External, non- accidental conditions (excluding mechanical loading), separately or in combination, that can be expected in service and which may affect the structure (e.g., temperature, moisture, UV radiation, and fuel).
(25) Fatigue or Environmental Damage. Structural damage related to fatigue or environmental effects such as delaminations, disbonds, splintering, or cracking.
(25) Fiber. A single homogeneous strand of material, essentially one-dimensional in the macro-behavior sense, used as a principal constituent in advanced composites because of its high axial strength and modulus.
(26) Fiber Volume. The volume of fiber present in the composite. This is usually expressed as a percentage volume fraction or weight fraction of the composite.
(28) Fill. The 90° yarns in a fabric, also called the woof or weft.
(29) Glass Transition. The reversible change in an amorphous polymer or in amorphous regions of a partially crystalline polymer from (or to) a viscous or rubbery condition to (or from) a hard and relatively brittle one.
(30) Glass Transition Temperature. The approximate midpoint of the temperature range over which the glass transition takes place.
(31) Hybrid. Any mixture of fiber types (e.g., graphite and glass).
(32) Impregnate. An application of resin onto fibers or fabrics by several processes: hot melt, solution coat, or hand lay-up.
(33) Intrinsic or discrete manufacturing defects. Intrinsic or discrete imperfections or flaws related to manufacturing operations, processing or assembly, such as voids, gaps, porosity, inclusions, fiber dislocation, disbonds, and delaminations.
(34) Lamina. A single ply or layer in a laminate in which all fibers have the same fiber orientation.
(35) Laminate. A product made by bonding together two or more layers or laminae of material or materials.
Page C - 136
(36) 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.
(37) Material System. The combination of single constituents chosen (e.g., fiber and resin).
(38) Material System Constituent. A single constituent (ingredient) chosen for a material system (e.g., a fiber, a resin).
(39) Matrix. The essentially homogeneous material in which the fibers or filaments of a composite are embedded in resins, which are mainly thermoset polymers in aircraft structure.
(40) 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.”
(41) 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.
(42) 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.
(43) Point Design. An element or detail of a specific design, which 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.
(44) 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.
(45) 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. Prepreg is usually drapable, tacky, and can be easily handled.
(46) 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.
Page C - 137
AC 29-2C, Chg 4 12/1/2011
(47) Residual Strength. The strength retained for some period of unrepaired use after a failure or partial failure due to fatigue, accidental, or discrete source of damage.
(48) Resin. An organic material with indefinite and usually high molecular weight and no sharp melting point.
(49) Resin Content. The amount of matrix present in a composite by either percent weight or percent volume.
(50) 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 COCURE).
(51) 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 remain suitable for its intended function.
(52) 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). Strain level is generally measured in thousandths of an inch per unit inch of part or microinches/inch (e.g., .003 in/in equals 3000 microinches/inch).
(53) Subcomponent. A major three- 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).
(54) Symmetrical Laminate. A composite laminate in which the ply orientation is symmetrical about the laminate midplane.
(55) Tape. Hot melt impregnated fibers forming unidirectional pre- preg.
(56) 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.
(57) 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).
(58) Warp. Yarns extended along the length of the fabric (in the 0° direction) and being crossed by the fill yarns (90° fibers).
Page C - 138
(59) Work Life. The period during which a compound, after mixing with a catalyst, solvent, or other compounding constituents, remains suitable for its intended use.
e. 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, MG 11 “Fatigue Evaluation of Rotorcraft Structure” AC 20-107 “Composite Aircraft Structure” AC 21-26 “Quality Control for the Manufacture of Composite Materials” CMH-17 “Composite Materials Handbook” AC 29-2C, MG 11 “Fatigue Tolerance Evaluation of Transport Category Rotorcraft Metallic Structure” DOT/FAA/CT-86/39 Whitehead, R.S., Kan, H.P., Cordero, R., and Seather, R., “Certification Testing Methodology for Composite Structures”, October 1986.
f. 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: fabrication requirements;
basic constituent, pre-preg and laminate material acceptance requirements, and material property determination requirements; protection of structure; lightning protection; static strength evaluation; damage tolerance and fatigue evaluation;
dynamic loading and response evaluation; and special repair and continued airworthiness requirements. Original as well as alternate or substitute material system constituents (e.g., fibers, resins), material systems (combinations of constituents and adhesives), and composite designs (e.g., laminates, cocured assemblies, bonded assemblies) 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 recommended that a FAA/AUTHORITY certification team approach be used for composite structural substantiation. The team should consist of FAA/AUTHORITY and cognizant members of the applicant’s organization. Personnel who are composites speciali sts (or are otherwise 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. The team should assure that permanent documentation 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), Page C - 139
AC 29-2C, Chg 4 12/1/2011 manufacturing processes and quality control, and Instructions for Continued Airworthiness (maintenance, overhaul, and repair manuals). The Airworthiness Limitations Section of the Instructions for Continued Airworthiness is approved by FAA engineering. Engineering practices for many of the areas identified below are available in CMH-17.
(1) The first area is the fabrication requirements of § 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. Some important concepts of AC 21-26 are included below.
(ii) Specific allowable defect limits (e.g., fiber waviness, warp defects, fill defects, porosity, hole edge effects, edge defects, resin content, large area disbonds, and delaminations) 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 created, validated, and approved for defect detection. Each critical engineering design should consider the variability of the manufacturing process to determine the worse case effects (maximum wavines s, 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 - see paragraph f.(6)) 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 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 certification team members (both the manufacturing inspection district office (MIDO) and FAA 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 CMH-17. These methods can also be used periodically for qualification status renewals of existing material systems and suppliers.
Page C - 140
(B) Receiving Inspection and In-Process Inspection. Applicants should be able to demonstrate to FAA certification team members (both MIDO and engineering), at any time, that their receiving and in-process quality control systems provide products, which continuously meet approved material and process specifications. Quality systems should be designed with appropriate checks and balances, so 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 witnessed.
(C) Material System Component Storage and Handling. Applicants should be able to demonstrate to FAA 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, so that the necessary statistical reliability and confidence levels for the items being inspected (that 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) 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 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 specifications should be approved and must comply with § 29.605. Any alternate specifications should provide at least the same level of quality and safety as the original specification. Any changes should be presented for FAA approval well in advance of the effective date of the production change.
(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 Page C - 141