MIL-HDBK-17B, Volume I, Section 2.3.2. These methods can also be used, periodically for qualification status renewals of existing material systems and suppliers.
(B) Receiving Inspection and In Process Inspection. Applicants should be able to demonstrate to FAA/AUTHORITY certification team members (both MIDO and engineering), at any time, that their receiving and in-process quality control systems provides products which continuously meet approved material and process specifications. Quality systems should be designed with appropriate checks and balances, such that the necessary statistical reliability and confidence levels for the items being inspected (that are specified by engineering) are continuously maintained.
This will require periodic standard inspections and engineering characterization tests on basic constituent and material system samples which should be conducted, as a minimum, on a batch-to-batch basis. The periodic testing necessary to maintain the quality standard should be conducted by the applicants on conformed samples and should be FAA/AUTHORITY-witnessed.
(C) Material System Component Storage and Handling. Applicants should be able to demonstrate to FAA/AUTHORITY certification team members (both MIDO and engineering), at any time, that their composite material system (or constituent) storage and handling procedures and specifications provide products which continuously meet approved material and process specifications. Quality systems should be designed with appropriate checks and balances, such that the necessary statistical reliability and confidence levels for the items being inspected (which are specified by engineering) are continuously maintained. This should require, as a minimum, periodic inspections to ensure that proper records are kept on critical parameters (e.g., room temperature "bench" exposure, shelf life, etc.) and that periodic basic constituent and material system characterization tests are conducted, on a batch-to-batch basis. The periodic testing necessary to maintain the quality standard should be conducted by the applicants on conformed samples and should be FAA/AUTHORITY-witnessed.
(D) Statistical Validation Level. It is necessary to maintain the minimum required statistical validation level of the quality control system (which should be specified for each critical item or constituent by the approved quality and engineering specifications). The statistical validation level should be defined and approved early in certification. Also, approval and proper usage should be continuously maintained during the entire procurement and manufacturing cycles.
(v) Alternate fabrication and process techniques should be approved and should comply with § 29.605. Any alternate techniques should provide at least the same level of quality and safety as the original technique. Any changes should be presented and FAA/AUTHORITY-approved well in advance of the change's production effectivity.
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(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" & "B" basis allowable strength values and other basic material properties (based on MIL-HDBK-17, or equivalent) are typically determined by small scale tests, such as coupon tests, for use in certification work. In the case of composites, determination of these basic constituent and material system properties will almost invariably involve the submittal, acceptance and use of company standards.
This is currently necessary because MIL-HDBK-17 has not completed development of "B" basis allowables for inclusion in the handbook. Also, test methods vary somewhat from manufacturer to manufacturer; therefore, individual company results will exhibit some scatter in final material property values. Any company standard which is approved and used should meet or exceed related MIL-HDBK-17 requirements.
Material structural acceptance criteria and property determination should, as a minimum, include the following:
(i) Property characterization requirements of all material systems (e.g., pre-pregs, adhesives, etc.) and constituents (e.g., fibers, resins, etc.) should be identified, documented, and approved. These requirements, once approved, should be placed in all appropriate procedures and specifications (such as those in Paragraph (g) (1) above).
(il) Moisture conditioning of test coupons, parts, subassemblies, or assemblies should be accomplished in accordance with MIL-HDBK-17, other similar approved methods or per FAA/AUTHORITY approved programs.
(iii) The maximum and minimum temperatures expected in service (as derived from test measurements, thermal analyses on panels and other parts, experience, or a combination) should be determined and accounted for in static and fatigue strength (including damage tolerance) substantiation programs considering associated humidity induced effects.
(iv) The glass transition temperature, Tg, is an important characteristic parameter of amorphous polymers, such as epoxies. It is the temperature below which the polymer behaves like a "glassy" solid and above which it behaves like a "rubbery" solid, i.e., it is the temperature at which there is a very rapid change in physical properties. In actuality, the change from a hard polymeric material to a rubbery material takes place over a narrow temperature range. A composite material will experience a drastic reduction in matrix controlled mechanical material properties when loaded in this temperature range. Since the resin (matrix) is the critical structural constituent in a composite and since Tg exceedance is critical to structural integrity; Tg determination is 1270 Par 788
necessary. The Tg margin methodology of MIL-HDBK-17, Section 2.2.2.1, should be implemented, i.e., the wet glass transition temperature (Tg) should be 50° F higher than the maximum structural temperature (see definition). For any type of resin or adhesive, an acceptable temperature margin using MIL-HDBK-17 techniques (e.g., consideration of limited high temperature excursions) or equivalent methodologies based on tests and/or experience should be established and approved early in the certification process. In no case should structural strength be degraded below limit load capability on a maximum world wide high temperature day.
(V) Local design values should be established by analysis and characterization tests and approved for specific structural configurations (point designs) which include the effects of stress risers (e.g., holes, notches, etc.) and structural discontinuities (e.g., joints, splices, etc.). Proper determination of these values for full-scale design and test should be considered one of the most critical building blocks in substantiating and evaluating a composite structure. These transitional load transfer areas typically produce the highest stresses (and strains) and serve as the nucleation sites for many of the failures (including those due to the relatively low interlaminar strength of composites) that occur in service in a full-scale part or assembly. Small scales tests (such as coupon, element, and subcomponent tests), or equivalent approved testing programs, and analytical techniques should be carefully designed, prepared, and approved to evaluate potential "hot spots" and provide accurate simulations and representations of full-scale article stresses and strains in the critical transition areas. Proper certification work in this area will ensure initial safety and continued airworthiness in full-scale production articles.
(vi) The design strain level for each major component and material system should be established and approved such that specified impact damage considerations are defined and properly limited. The effects of the approved strain levels should be established for each composite material using small scale characterization tests and the results should be used to establish or verify the maximum allowable design strain level for each full-scale article. The maximum allowable design strain values selected should also take into account the reliability and confidence levels established for the relevant portions of the quality control system. This methodology is necessary because the amount and size of flaws in the production article may restrict the allowable level of design strain. In a no-flaw-growth design, the maximum specified impact damage and manufacturing flaw size at the most critical location on the part will be a major factor in determining the maximum allowable elastic strain. This design approach is currently selected for nearly all civil and most military applications; since, under normal conditions, only visual inspections are required in the field (unless unusual external damage circumstances such as a hail storm occur) to maintain the initial level of airworthiness (safety). However, many military applications because of their demanding missions, employ scheduled field non-destructive inspection (NDI) maintenance, (such as comparative ultrasonics) to ensure that flaw growth either does not occur, is controlled by approved structural repair, or by replacement of affected parts. To date, civil applications have not been presented that desire a flaw growth, Par 788 1271
phased NDI approach. Therefore, selection of the full-scale article's design strain limit based on small scale tests for a no flaw growth design is seen to be extremely important.
(vii) Composite and adhesive properties should be determined such that detrimental structural creep does not occur under the sustained loads and environments expected in service. Small scale characterization tests (such as coupon, element, and subcomponent tests) and analysis, which verify and establish the full-scale design criteria and parameters necessary to ensure that detrimental structural creep in full-scale structure does not occur in service, should be conducted early in certification and should be FAA/AUTHORITY-approved.
(viii) Material allowable strength values for full-scale design and testing should be developed using the coupon procedures presented in MIL-HDBK-17 or equivalent. At least three batches of material samples should be used in material allowable strength testing. Company standards should be prepared, evaluated and FAA/AUTHORITY approved early in certification (as part of the building block process), that reflect the material property determination considerations recommended in MIL-HDBK-17 on a equal to or better than basis.
(3) The third area is the protection of structure as required by § 29.609.
Protection against thermal and humidity effects and other environmental effects (e.g., weathering, abrasion, fretting, hall, ultraviolet radiation, chemical effects, accidental damage, etc.) should be provided, or the structural substantiation should consider the results of those effects for which total protection is impractical. Determination and approval of worst-case or most conservative operating limits, and damage scenarios should be accomplished. Appropriate flammability and fire resistance requirements should also be considered in selecting and protecting composite structure. Usually a hazard analysis is conducted early in certification which identifies the various threats and threat levels for which protection must be provided. This data is then used to construct and submit for approval the methods-of-compliance necessary to provide proper structural protection.
(4) The fourth area is the lightning protection requirements of § 29.610.
Protection should be provided and substantiated in accordance with analysis and with tests such as those of AC 20-53A and FAA Report DOT/FAA/CT-86/8. For composite structure projects involving rotorcraft certified to earlier certification bases (which do not automatically include the lightning protection requirements of § 29.610), these requirements should be imposed as special conditions. The design should be reviewed early in certification to ensure proper protection is present. The substantiation test program should also be established, reviewed and approved early to ensure proper substantiation.
(5) The fifth area is the static strength evaluation requirements of §§ 29.305 and 29.307 for composite structure. Only conservative proven methods of static 1272 Par 788
analysis and failure criteria should be employed. The material stress-strain curve should be clearly established, at least through the ultimate design load, for each composite design. Composite structure should be statistically demonstrated, incrementally, through a program of analysis, coupon tests, minor component ultimate load tests and major component ultimate load tests. The static strength substantiation program should consider all critical loading conditions for all critical structure including residual strength and stiffness requirements after a predetermined length of service,
e.g., end of life (EOL) (which takes into account damage and other degradation due to the service period). Analytical reports and tests should consider all possible failure modes and should include the critical, allowable effects of:
Environment (reference Paragraphs 2 and 3.)
(ii) Service Life (residual limit strength and stiffness demonstration.)
(iii) Load path loss (fail-safe analysis and limit strength demonstration.)
(iv) The standard fabrication process and its variability.
(v) Impact damage expected during service up to the established threshold of detectability of the field inspection methods to be employed.
(vi) Point design and structural discontinuity considerations (e.g., stress risers, joints, etc.).
(vii) Unless the ultimate strength of each critical bonded joint can be reliably substantiated in production by NDI techniques (or other equivalent, approved techniques), then limit load capability must be guaranteed by either of the following or a combination thereof:
(A) The maximum disbond of each critical bonded joint which will carry limit load is established by test, analysis, or both. Disbonds greater than these values are typically prevented by design features.
(B) Each critical bonded joint on each production article should be proof tested to the critical limit load.
(vill) For static strength analysis laminae and laminate "A" and "B" basis allowables (determined in accordance with Paragraph (2)) should be used subject to the following conditions unless lower material properties are required by point design considerations (e.g., stress risers, joints, etc.) stiffness requirements (e.g., flutter or vibration margins), fatigue strength (including damage tolerance), or other overriding considerations.
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(A) When applied loads are distributed through a single load path or single member within an assembly, the failure of which would result in the loss of the structural integrity of the component involved or inability of the rotorcraft structure to carry limit load, the part should be designed, analyzed, and tested using "A" basis allowables.
(B) Redundant (fall-sate) structures in which the failure of individual elements would result in applied loads being safely redistributed to other load carrying members without exceeding the limit load capability of the rotorcraft structure may be designed, analyzed, and tested using "B" basis allowables.
(6) The sixth area is the fatigue evaluation requirements of § 29.571. The fatigue evaluation method for the rotorcraft being certified should consider damage tolerance in accordance with AC 20-107A.
(i) The safe-life method for composite structure as defined in AC 20-107A is a flaw tolerant safe-life method (e.g., the test specimens consider inherent production flaws and impact damage (reference Paragraph (7)i)).
(її) Large area disbonds, weak bonds, delaminations, or other defects should be considered in tests or be prevented or be limited by appropriate flaw tolerant special design features and by special manufacturing, maintenance, and inspection procedures. Special attention should be assigned to all pure bond lines (reference Paragraph (5)).
(ill) Non-fail-safe or partially fail-safe dynamic component structure, which may employ bond lines as the only load path, should be designed to relatively small previously approved values of elastic, ultimate strain for the material system utilized, and should be subjected to full-scale S-N curve testing. Six or more specimens are recommended, as part of the substantiation process. Where practical, flight-by-flight spectrum testing should be used.
(iv) All critical safety of flight composite structure must be designed to be flaw (damage) tolerant. Environment degradation and in-service damage critical values are typically included in the flaw tolerance evaluation. All other key factors, such as material selection, manufacturing, and quality assurance controls, and in-service inspection and maintenance, as noted previously, are also to be accounted for.
(V) The fail-safe design features of the rotor heads and blade retention systems, other critical primary composite structure, and point design features (e.g., bonded metal-to-composite joints) should be assessed and appropriate inspection programs provided to prevent catastrophic failure from flaw/damage propagation.
(vi) The method of generating S-N curves using approved raw data should be demonstrated, evaluated, and approved.
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(vii) Any limited life items must be identified and placed in the Airworthiness Limitations section of the maintenance manual in accordance with § 29.571.
(viii) Load spectra, load truncation methods and all other major aspects of the fatigue evaluation are documented in test proposals and approved.
(ix) Flaw growth rates (from initial detectability to the established value for residual strength) must be previously established and closely monitored during substantiation. This data should be used to establish special phased inspections and maintenance intervals for critical structure, as required.
(7) The seventh major area is the dynamic loading and response requirements of § 29.629 for vibration and resonance frequency determination and separation for aeroelastic stability and stability margin determination for flutter critical flight structure.
Critical parts, locations, excitation modes, and separations are to be identified and substantiated. This substantiation should consist of analysis supported by tests and tests which account for repeated loading effects and environment exposure effects on critical properties, such as stiffness, mass, and damping. Initial stiffness, residual stiffness, proper critical frequency design, and structural damping are provided as necessary to prevent vibration, resonance, and flutter problems.
All vibration and resonance critical composite structure are identified and properly substantiated.
All flutter-critical composite structure are identified and properly substantiated. This structure must be shown by analysis to be flutter free to 1.1 VNE (or any other critical operating limit, such as Vp, for a VSTOL aircraft) with the extent of damage for which residual strength and stiffness are demonstrated.
(ili) Where appropriate, crash impact dynamics considerations should be taken into account to ensure proper crash resistance and a proper level of occupant safety for an otherwise survivable impact.
(8) The eighth area is the special repair and continued airworthiness requirements of §§ 29.611, 29.1529, and FAR Part 29 Appendix A for composite structures. When repair and continued airworthiness procedures are provided in service documents (including approved sections of the maintenance manual or instructions for continued airworthiness) the resulting repairs and maintenance provisions must be shown to provide structure which continually meets the guidance of Paragraphs (1) through (7) of this AC paragraph. All certification based repair and continued airworthiness standards, limits, and inspections must be clearly stated and their provisions and limitations defined and documented to ensure continued airworthiness. In general, no composite repair should be attempted which is out of Par 788 1275
scope to repairs stated in an approved Structural Repair Manual (SRM) without an engineering design approval by a qualified FAA/AUTHORITY representative (DER or staff engineer). The following minimum criteria should be met in any acceptable composite repair:
(i) The repair should be permanent.
(i) The repair should restore the structure to the required strength and stiffness.
The repair should restore all functional requirements.
(iv) The repair should have a negligible weight penalty.
(v) The repair should be aerodynamically compatible.
(vi) The repair materials should be compatible in all essential aspects with the parent materials.
In summary, primary composite structure is an especially critical structure that requires a clearly defined, phased approval (building block) certification process. This process should involve the entire project certification team from a project's start to its finish so that proper certification is continuously and ultimately achieved. Also, in some special cases, involving new advanced state-of-the-art composite technology, an issue paper may be necessary. However, in the majority of cases (using current composite materials and design philosophy) the applicant's acknowledged use of this advisory material (as recorded in the type board minutes) should eliminate the need for a separate issue paper.
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Appendix 1 APPENDIX 1
FATIGUE EVALUATION OF TRANSPORT CATEGORY
ROTORCRAFT STRUCTURE (INCLUDING FLAW TOLERANCE) PURPOSE. This advisory material provides an acceptable means of compliance with the provisions of § 29.571 of the Federal Aviation Regulations (FAR) dealing with the fatigue evaluation of transport category rotorcraft structure (AC 20-95, Fatigue Evaluation of Rotorcraft Structure, May 18, 1976, applies to normal category rotorcraft structure and older transport category rotorcraft). The fatigue evaluation procedures outlined in this Appendix are for guidance purposes only and are neither mandatory nor regulatory in nature. Although a uniform approach to fatigue evaluation is desirable, it is recognized that in such a complex problem, new design features and methods of fabrication, new approaches to fatigue evaluation, and new configurations may require variations and deviations from the procedures described herein. It is recommended that major deviations from the procedures be coordinated with the Rotorcraft Standards Staff, ASW-110, to assure national standardization.
2. SPECIAL CONSIDERATIONS. The structure of rotorcraft is subject to cyclic stresses in practically every regime of flight. In addition, since rotorcraft are highly maneuverable and capable of forward, rearward, sideward, vertical, and rotational flight, operating limitations due to fatigue are possible in practically all flight situations.
Corrosion and other environmental damage are also common in rotorcraft operations.
For these reasons, special attention should be focused on the fatigue evaluation of rotorcraft structure.
3. BACKGROUND.
a. During recent years there have been significant state-of-the-art and industry practice developments in the area of structural fatigue and fail-safe strength evaluations of transport category rotorcraft. The advance in the state-of-the-art has resulted from two primary programs: (1) the perfecting of production techniques for composite construction, and (2) the damage tolerant design features required to meet the battle damage requirements of military programs such as the Advanced Attack and Heavy Lift Helicopter (AAH and HLH) programs.
b. Recognizing that advances in state-of-the-art and industry practice warranted changes to the existing fatigue requirements in Part 29, the regulatory requirements of § 29.571 were substantially revised. The revision to § 29.571 requires new guidance material containing compliance provisions related to the changes. Also, this Appendix supplements AC 20-95 for new transport category rotorcraft and provides guidance material in the flaw tolerance area. General guidance material for flaw/damage tolerance of composite structures is provided in AC 20-107A, Composite Aircraft Structure, April 25, 1984.
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Appendix 1
4. INTRODUCTION
a. Definitions.
(1) Fatigue Tolerance. The capability of structure to continue functioning without catastrophic failure after being subjected to fatigue (repeated) loads expected during operation of the rotorcraft. Fatigue tolerance may be achieved by safe-life design, flaw tolerant safe-life design (enhanced safe life), fail-safe design considering flaw growth, or a combination.
(i) Safe Life. The capability of pristine structure as shown by tests, or analysis based on tests, not to sustain measurable cracks during the service life of the rotorcraft or before an established replacement time. Special inspection intervals or other special procedures are not usually prescribed for safe-life substantiations.
(Routine inspections for wear, fretting, corrosion, crack, and service damage as outlined in § 29.1529 are, of course, appropriate.)
(ii) Flaw Tolerant (Enhanced) Safe Life. The capability of flawed structure as shown by tests or analysis based on tests to sustain, without measurable flaw growth, the spectrum of operating loads expected during the service life of the rotorcraft or during an established replacement time. Measurable flaw growth means flaw growth beyond an acceptable threshold of detectability. Enhanced safe life may be achieved by designing structure (single element or multiple element) that provides for resistance to crack initiation from manufacturing or service flaws, by material selection, by material processing, by limitation of stress levels, and by geometric design features. The protection of structure against environmental damage and accidental mechanical damage through the use of protective coatings may be used in determining initial flaw types and sizes to be considered."
(III) Fall-Sate Design Considering Flaw Growth. The capability of rotorcraft structure to continue functioning without catastrophic failure after being subjected to fatigue damage, corrosion, intrinsic flaws, or accidental damage expected during fabrication and operation of the rotorcraft. The concept of fail-safety now includes some discrete damage requirements and more explicitly requires consideration of "life-remaining" after flaws occur until the flaws are detected using a prescribed inspection plan or until the part is replaced. Fail-safe designs may include any of the following features:
(A) 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.
(1) Active Multiple Load Path. Structure providing two or more load paths that are all loaded during operation to a similar load spectrum. The use of active Page A1- 2
Appendix 1 multiple load paths requires special attention to fatigue damage in the remaining members after failure of a member.
(2) 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.
(B) Flaw Arrest (Flaw Stopper) Feature. Structure that does not provide completely separate and distinct load paths but does provide features of design such as bonded and/or riveted straps, changes in geometry, or special processing techniques such as rolling or coining to retard or arrest flaw growth.
(C) Slow Flaw Growth Feature. Structure (single element or multiple element) that provides for slow flaw growth by material selection, material processing, limitation of stress levels, geometrical design features, or by other methods.
NOTE: The detection of the crack or other flaw is an integral part of the flaw growth method. The need to use complicated inspection techniques may not be practical in some cases. See the discussion in Paragraph (7)(a)(3)(ii).
(2) Flaw. Structural imperfections in excess of type design allowances for "as manufactured" or "pristine" structure.
(1) In Metallics.
(A) For Fail-Safe Crack Growth. Corner cracks for holes, semicircular cracks for surfaces, realistic cracks for other locations.
(B) For Flaw Tolerance (Enhanced) Safe Life. Gouges, scratches, corrosion, fretting, or wearing likely to occur during fabrication and operation of the rotorcraft.
(il) In Nonmetallics. Flaws should be determined in accordance with AC 20-107A guidance.
(3) Limit Design Load. "The maximum loads to be expected in service," as defined by § 29.301(a), are considered as limit loads for new structure and as ultimate loads for flaw tolerance residual strength demonstration purposes. The residual strength after failure should equal or exceed these loads for flaw tolerance.
(4) Damage Tolerance. No definition is provided in this Appendix for this expression because many definitions have been applied by the civil and military communities. The expression has been used to describe design features, structural systems (including design, manufacturing, and operating considerations), and substantiation techniques such as fracture mechanics analyses for metallics. This Page A1-3
Appendix 1 Appendix uses expressions for those features of the general damage tolerance concept used to provide and demonstrate fatigue tolerance of rotorcraft structure including consideration of flaws. Tolerance to flaws rather than tolerance to damage is specified in this Appendix since flaws intrinsic to certain manufacturing and fabrication processes are covered in addition to damage from handling, corrosion, etc.
b. Rotorcraft Flaw Tolerance. Flaw tolerant design as substantiated by fail-safe flaw growth or flaw tolerant (enhanced) safe-life means outlined in § 29.571 and Paragraph 7 of this Appendix is required, unless it entails such complications that an effective flaw tolerant structure cannot be achieved within the limitations of geometry, inspectability, or good design practice. Good design practice includes consideration of component complexity, component weight, methods of production and component cost.
Under these circumstances, a design that complies with safe-life criteria should be used. Typical examples of structure that might not be conducive to flaw tolerance design are swashplates, main rotor shafts, push rods, small rotor head components
(l.e., devices, bolts, etc.), landing gear, and gearbox internal parts including bearings.
In addition, the need for the use of inspection techniques and equipment or highly trained personnel--resources not available (for economic or other reasons) to the small operator or in remote areas of operation--should be carefully considered (reference Paragraphs 4a(1)(iii)(C) and 7a(3)i)).
c. Test Background. Experience with the application of methods of fatigue evaluation indicates that a relevant test background should exist in order to achieve the design objective. Even under the flaw tolerance method discussed in Paragraph 7, it is the general practice within industry to conduct flaw tolerance tests for design information and guidance purposes. Flaw location and flaw growth data based on test results and service history of similar parts, if available, should also be considered in establishing a recommended inspection program.
d. Manufacturing Considerations. Assurance of structural adequacy also includes manufacturing and fabrication in accordance with design requirements and specifications, quality control to monitor compliance, and effective service inspection procedures.
e. Fatigue Tolerance Considerations. In the fatigue tolerance evaluation, the following items should be considered:
(1) Identification of the structure to be considered in each evaluation (a failure mode and effects analysis or similar method should be used).
(2) The stresses and strains (steady and oscillatory) associated with all representative steady and maneuvering operating conditions expected in service.
(3) The frequency of occurrences of various flight conditions and the corresponding spectrum of loadings and stresses.
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