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 § 27.605:
(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) 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 Page MG 8 - 9
AC 27-1B, Chg 2 4/25/06 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.
(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.
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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 § 27.605. 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 §§ 27.603 and 27.613. 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 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 Page MG 8 - 11
AC 27-1B, Chg 2 4/25/06 the polymer behaves like a “glassy” solid and above which it behaves like a “rubbery” solid, i.e., it is the temperature at which there is a very rapid change in physical properties. In actuality, the change from a hard polymeric material to a rubbery material takes place over a narrow temperature range. A composite material will experience a drastic reduction in matrix controlled mechanical material properties when loaded in this temperature range. Since the resin (matrix) is the critical structural constituent in a composite and since Tg exceedance is critical to structural integrity, Tg determination is necessary. The Tg margin methodology of MIL-HDBK-17 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 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).
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However, many military applications, because of their demanding missions, employ scheduled field non-destructive inspection (NDI) maintenance, (such as comparative ultrasonics) to ensure that flaw growth either does not occur, is controlled by approved structural repair, or by replacement of affected parts. To date, civil applications have not been presented that desire a flaw growth, phased NDI approach. Therefore, selection of the full-scale article’s design strain limit based on small-scale tests for a no flaw growth design is seen to be extremely important.
(vii) Composite and adhesive properties should be determined such that detrimental structural creep does not occur under the sustained loads and environments expected in service. Small-scale characterization tests (such as coupon, element, and subcomponent tests) and analysis, which verify and establish the full-scale 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. The intent is to represent the material variability including the effects that can occur in multiple batches of material and process runs. 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 § 27.609.
Protection against thermal and humidity effects and other environmental effects (e.g., weathering, abrasion, fretting, hail, ultraviolet radiation, chemical effects, accidental damage, etc.) should be provided, or the structural substantiation should consider the results of those effects for which total protection is impractical. Determination and approval of worst-case or most conservative operating limits, and damage scenarios should be accomplished. Appropriate flammability and fire resistance requirements should also be considered in selecting and protecting composite structure. Usually a threat analysis is conducted early in certification that 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 § 27.610.
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 § 27.610), 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 Page MG 8 - 13
AC 27-1B, Chg 4 3/31/2014 program should also be established, reviewed, and approved early to ensure proper substantiation.
(5) The fifth area is the static strength evaluation requirements of §§ 27.305 and 27.307 for composite structure. Structural static strength substantiation of a composite design should consider all critical load cases and associated failure modes, including effects of environment, material and process variability, and defects or service damage that are not detectable or allowed by the quality control, manufacturing acceptance criteria, or maintenance documents of the end product. The static strength demonstration should include a program of component ultimate load tests, unless experience exists to demonstrate the adequacy of the analysis, supported by subcomponent tests or component tests to accepted lower load levels. The necessary experience to validate an analysis should include previous component ultimate load tests with similar designs, material systems, and load cases.
(i) The effects of repeated loading and environmental exposure, both of which may result in material property degradation, should be addressed in the static strength evaluation. This can be shown by analysis supported by test evidence, by tests at the coupon, element or subcomponent levels, or alternatively by existing data.
Earlier discussions in this AC address the effects of environment on material properties (reference paragraph g(2) of this AC paragraph) and protection of structure (reference paragraph g(3) of this AC parapraph). Static strength tests should be conducted for substantiation of new structure. For the critical loading conditions, two approaches to account for prior repeated loading and environmental exposure for structural substantiation exist.
• In the first approach, the large-scale static test should be conducted on structure with prior repeated loading and conditioned to simulate the environmental exposure and then tested in that environment.
• The second approach relies upon coupon, element, and sub-component test data to assess the possible degradation of static strength after application of repeated loading and environmental exposure. The degradation characterized by these tests should then be accounted for in the static strength demonstration test (e.g., load enhancement), or in the analysis of these results (e.g., showing a positive margin of safety with allowables that include the degrading effects of environment and repeated load).
In practice, the two approaches may be combined to get the desired result (e.g., a large-scale static test may be performed at temperature with a load enhancement factor to account for moisture absorbed over the aircraft structure’s life).
(ii) The strength of the composite structure should be statistically established, incrementally, through a program of analysis and tests at the coupon, element, subcomponent, or component levels. As part of the evaluation, building block tests and analyses at the coupon, element, or subcomponent levels can be used to Page MG 8 - 14
address the issues of variability, environment, structural discontinuity (e.g., joints, cutouts or other stress risers), damage, manufacturing defects, and design or processspecific details. Figure AC 27 MG 8-1 provides a conceptual schematic of tests included in the building block approach. The material stress-strain curve should be clearly established, at least through the ultimate design load, for each composite design. As shown in Figure AC 27 MG 8-1, the large quantity of tests needed to provide a statistical basis comes from the lowest levels (coupons and elements) and the performance of structural details are validated in a lesser number of sub-component and component tests. The static strength substantiation program should also 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).
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AC 27-1B, Chg 2 4/25/06
COMPONENTS
TAIL ROTOR BLADE
NON-GENERIC SPECIMENS SUB-COMPONENTS
DETAILS STRUCTURAL FEATURES
SHI REA
ELEMENTS SKIN + WEB
GENERIC SPECIMENS SK
COUPONS DATA BASE Figure AC 27 MG 8-1: Schematic diagram of building block tests.
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(iii) Allowables should be used as specified in § 27.613. These allowables may be generated at the lamina, laminate, or specific design feature level
(e.g. filled hole, lap joint, stringer run-out, etc.), provided they accurately reflect the actual value and variability of the structural strength for the critical failure modes being considered, at each point design where margins need to be established.
(iv) The static test articles should be fabricated and assembled in accordance with production specifications and processes so that they are representative of production structure including defects consistent with the limits established by manufacturing acceptance criteria.
(v) The material and processing variability of the composite structure should be considered in the static strength substantiation. This can be achieved by establishing sufficient process and quality controls to manufacture structure and reliably substantiate the required strength in tests and analyses, which support a building block approach. If sufficient process and quality controls cannot be achieved, it may be necessary to account for greater variability with special factors (§ 27.619) applied to the design. Such factors should be accounted for in the component static tests or analysis.
(vi) It should be shown that impact damage (or other minor discrete source damage) that can be realistically expected from manufacturing and service, but not more than the established threshold of detectability for the selected inspection procedure, will not reduce the structural strength below ultimate load capability. This static strength capability can be shown by analysis supported by test evidence, or by a combination of tests at the coupon, element, subcomponent, and component levels.
Later discussions in this AC paragraph address the issues associated with damage in excess of that considered in g(5) of this AC paragraph and drops in residual strength below ultimate load capability (reference paragraph g(6)) below.
(6) The sixth area is the fatigue evaluation of structure requirements of § 27.571.
(i) FATIGUE EVALUATION - BACKGROUND. The static strength determination required by §§ 27.305 and 27.307 establishes the ultimate load capability for composite structures that are manufactured, operated, and maintained with established procedures and conditions. The fatigue tolerance evaluation required by § 27.571 establishes procedures that allow the composite structure to retain the intended ultimate load capability when subjected to expected fatigue loads and conditions during its operational life. The procedures established by the fatigue tolerance evaluation include component retirement times and/or inspection intervals.
The fatigue tolerance evaluation requires a flaw tolerance assessment that assumes that the baseline ultimate strength capability might be compromised by damage caused by fatigue, environmental effects, intrinsic discrete flaws, or accidental damage. The flaw tolerance assessment establishes procedures that do not allow the static strength capability to degrade below the ultimate strength capability for extended periods, Page MG 8 - 17
AC 27-1B, Chg 2 4/25/06 assuming such damage occurs within the operational life of the structure. When this damage occurs, the remaining structure will withstand reasonable loads without failure or excessive structural deformations until the damage is detected and the component is either repaired to restore ultimate load capability or retired.
(ii) FLAW TOLERANCE EVALUATION - GENERAL. The nature and extent of the required analysis or tests on complete structures or portions of the primary structure can be based on applicable previous fatigue or damage tolerant designs, construction, tests, and service experience on similar structures. In the absence of experience with similar designs, FAA/AUTHORITY-approved structural development tests of components, subcomponents, and elements should be performed. The following considerations are unique to the use of composite material systems and should be observed for the method of substantiation selected by the applicant.
Rotorcraft structure provides a broad range of composite applications that are quite different in terms of functionality, geometry, and inspectability. These include the rotors, the drive shafts, the fuselage, control system components (e.g., push-pull rods), and the control surfaces. When selecting the approach, attention should be given to the composite application under evaluation, the type of potential damage or degradation of the structural design details, the materials used, and margin over flight loads. Whatever the approach that may be selected, the following considerations will apply for tests and analysis:
(A) The test articles should be fabricated and assembled in accordance with production specifications and processes so that the test articles are representative of production structure.
(B) The test articles should include material imperfections whose extent is not less than the limits established under the inspection and acceptance criteria used during the manufacturing process and consistent with the inspection techniques used in service (e.g., visual, ultrasonic, X-ray). The initial extent of these imperfections should be discussed and agreed with the FAA/AUTHORITY, taking into account experience in manufacturing and routine in-service inspections. Typical defects to be considered include but are not limited to the following:
(1) Disbonds and weak bonds (considered as disbonds).
(2) Delaminations, fiber waviness, porosity, voids.
(3) Scratches, gouges, and penetrations.
(4) Impact damage.
(C) The use of composite secondary bonding in manufacturing or maintenance requires strict process and quality controls to achieve the reliability needed to use such technology in critical structures (reference AC 21-26). Assuming good process and quality controls, service history has shown that additional damage tolerant Page MG 8 - 18
design considerations are also needed to ensure the safety of structure with secondary bonds (i.e., random, but an unacceptable numbers of weak bonds discovered in service). Unless the ultimate strength of each critical bonded joint can be reliably substantiated in production by NDI techniques (or other equivalent, approved techniques), then the limit load capability should be ensured by any of the following or a combination thereof.
(1) Consider isolated disbonds and weak bonds (represented by zero bond strength) in structural elements that use secondary bonding for primary load transfer. The associated disbond size should be up to the limitations provided by redundant design features (i.e., mechanical fasteners or a separate bonding detail).
The structure containing such damage should be shown to carry limit load by tests, analyses, or some combination of both. For purposes of test or analysis demonstration, each disbond should be considered separately as a random occurrence (i.e., it is not necessary to demonstrate residual strength with all structural elements disbonded simultaneously).
(2) Each critical bonded joint on each production article should be proof tested to the critical limit load.
(3) Critical bonded joints that have high static margins of safety (e.g., some rotor blades) may be acceptable, provided there is satisfactory service history of like or similar components.
(D) The fatigue load spectrum developed for fatigue testing and analysis purposes should be representative of the anticipated service usage. Low amplitude load levels that can be shown not to contribute to fatigue damage may be omitted (truncated). Reducing maximum load levels (clipping) is generally not accepted.
(E) Environmental effects (temperature and humidity representative of the expected service usage) on the static or fatigue behavior and damage growth should be considered. Unless tested in the environment, appropriate environmental knock down factors for the static and the fatigue test articles should be derived and applied in the evaluation.
(F) Variability in fatigue behavior should be covered by appropriate load and/or life scatter factors and these factors should take into account the number of specimens tested.
(G) The following Figure AC 27 MG 8-2 illustrates the extent of the impact damage that needs to be considered in the flaw tolerance evaluation.
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AC 27-1B, Chg 2 4/25/06 Figure AC 27 MG 8-2: Characterization of Impact Damage
(1) Both the energy level associated with the static strength demonstration and the maximum energy level associated with the damage tolerance evaluation (defined in the figure above) are dependent on the part of the structure under evaluation and a threat assessment.
(2) Obvious impact damage is used here to define the threshold from which damage is readily detectable and appropriate actions taken before the next flight.
(3) Barely Detectable Impact Damage defines the state of damage at the threshold of detectability for the approved inspection procedure. Barely Visible Impact Damage (BVID) is that threshold associated with a detailed visual inspection procedure.
(4) Detectable Damage defines the state of damage that can be reliably detected at scheduled inspection intervals. Visible Impact Damage (VID) is that state associated with a detailed visual inspection.
(5) A threat assessment is needed to identify impact damage severity and detectability for design and maintenance. A threat assessment usually includes damage data collected from service plus an impact survey. An impact survey consists of impact tests performed with configured structure, which is subjected to boundary conditions characteristic of the real structure. Many different impact scenarios and locations are typically considered in the survey, which has a goal of identifying the most Page MG 8 - 20