AC 27-1B, Chg 4 12/1/2011
(1) The first area is the fabrication requirements of § 27.605:
(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 detect limits (e.g., fiber waviness, warp detects, 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 worst case effects (maximum waviness, disbonds, delaminations, and other critical defects) allowed by the reliability limitations of the approved inspection program.
(ili) 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.
(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 Page C - 96
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 § 27.605. 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 §§ 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 material properties (based on CMH-17 or equivalent procedures) are typically Page C - 97
AC 27-1B, Chg 4 12/1/2011 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 the FAA (new managers of CMH-17) has not completed development of "B" basis allowables for inclusion in CMH-17. 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 used should meet or exceed related CMH-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) and constituents (e.g., fibers, resins) should be identified, documented, and approved. These requirements, once approved, should be placed in all appropriate procedures and specifications such as those in paragraph f.(1).
(ii) Moisture conditioning of test coupons, parts, subassemblies, or assemblies should be accomplished in accordance with CMH-17, other similar approved methods or per FAA approved programs.
(ili) 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 humidityinduced effects.
(iv) The wet 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 (l.e., it is the temperature at which there is a very rapid change in physical properties). 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 is the critical structural element in a composite matrix and the Ty is critical to structural integrity, a Ty determination is necessary. The To margin methodology of CMH-17 should be implemented (i.e., the 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 CMH-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) and structural discontinuities (e.g., joints, splices). Proper determination of these values for full-scale design and test should be considered one of the most critical building blocks in substantiating and Page C - 98
evaluating a composite structure. These transitional load transfer areas typically produce the highest stresses (and strains) and serve as the initiation 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 scale 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 so 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 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 applicants have not requested 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 extremely important.
(vil) Composite and adhesive properties should be determined so 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-approved.
(vili) Material allowable strength values for full-scale design and testing should be developed using the coupon procedures presented in CMH-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 Page C - 99
AC 27-1B, Chg 4 12/1/2011 samples should be used in material allowable strength testing. Company standards should be prepared, evaluated and FAA-approved early in certification (as part of the building block process), that reflect the material property determination considerations recommended in CMH-17 on an equal to or better than basis.
(3) The third area is the protection of structure as required by § 27.609. Protection against thermal, humidity, and other environmental effects (e.g., weathering, abrasion, fretting, hail, ultraviolet radiation, chemical effects, accidental damage) 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 the certification process 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 20-53, "Protection of Aircraft Fuel Systems Against Fuel Vapor Ignition Caused by Lightning" and FAA Report DOT/FAA/CT-86/8. For composite structure projects involving rotorcraft certificated 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 the certification process 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 §§ 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.
(1) 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 (see paragraph f.(2)) and protection of structure (see paragraph f.(3)). Static strength tests Page C - 100
should be conducted for substantiation of new structure. For the critical loading conditions, two approaches to account for prior repeated loading or 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 a 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 address the issues of variability, environment, structural discontinuity (e.g., joints, cut-outs or other stress risers), damage, manufacturing defects, and design or process-specific details.
Figure AC 27.573-1 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
27.573-1, 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 4 12/1/2011
COMPONENTS
TAIL ROTOR BLADE NON-GENERIC SPECIMENS •SUB-COMPONENTS
DETAILS STRUCTURAL FEATURES
ELEMENTS SKIN
• WEB GENERIC SPECIMENS SKIN •
COUPONS DATA BASE Figure AC 27.573-1. Schematic diagram of building block tests.
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(iii) Allowables should be evaluated and 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), 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 address the issues associated with damage in excess of that considered in f.(5) and drops in residual strength below ultimate load capability (see paragraph f.(6) below).
(6) The sixth area is the damage tolerance and fatigue evaluation requirements of § 27.573.
(i) 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 damage tolerance and fatigue evaluation required by § 27.573 mandates 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 requirements established for the damage tolerance and fatigue evaluation include component replacement times, inspection intervals, or other procedures as necessary to avoid catastrophic tailure. These evaluations assume that the baseline ultimate strength capability might be compromised by damage caused by fatigue, environmental effects, intrinsic or discrete flaws, or accidental damage. This damage includes flaws or defects, which may occur in manufacturing or maintenance and which are used to set the ultimate strength capability and establish the manufacturing acceptance criteria. The damage tolerance assessment establishes Page C - 103
AC 27-1B, Chg 4 12/1/2011 standards that allow the static strength capability to degrade below the ultimate strength capability assuming such damage occurs within the operational life of the structure.
However, when this damage occurs, the remaining structure must withstand expected loads without failure or excessive structural deformations until the damage is detected and the component is either repaired to restore ultimate strength capability or retired.
(ii) General. The nature and extent of the required analysis or tests on complete structures and 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 and degradation of the structural design details, the materials used and margin over flight loads. Whatever the approach 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, taking into account experience in manutacturing 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.
All of the damages identified in the preceding paragraph (B) above should be derived from the threat assessment described in the following paragraph (C).
(C) For each PSE, a threat assessment must be made of the probable locations, types, and sizes of damage considering fatigue, environmental effects, intrinsic and Page C - 104
discrete flaws, and impact or other accidental damage. This determination must be submitted with accompanying rationale to the FAA/AUTHORITY for approval. This rationale may include experience with similar materials, designs, processes (manufacturing, maintenance, and overhaul), structural details, or structure, and may also include service failure evaluations, manufacturing records, overhauls and repair reports, field service reports, incident and accident investigations, service impact surveys, inspectability surveys, and engineering judgment. Consideration should also be given to factors that:
- Reduce scatter and deviations from nominal structures, such as frozen processes, Flight Critical Parts programs, and materials and manufacturing processes to mitigate intrinsic flaws (inclusions and defects).
- Preclude a type of damage by use of a specific design feature (material selection, surface treatment, protective coating, or shielding), a specific stress level (for fatigue damage), or a specific manufacturing inspection process (if it can be shown to be highly reliable, wellcontrolled, documented, and systematically required).
The assessment should include:
• A systematic evaluation of all the location, types, and sizes of damage and their estimated probability of occurrence.
• A selection or elimination of this damage based on the above estimate.
• A verification that the inspection method selected is capable of detecting the damage at the size and location determined.
The types of damage to consider include:
(1) Intrinsic Flaws (imperfections), which are probable to exist in an asmanufactured structure based on the evaluation of the details and potential sensitivities of the specific manufacturing work processes used. The types of flaws to be considered include voids, disbonds, inclusions, foreign objects, resin-rich and resin-starved areas, and improper ply orientation or ply ending. The sizes of the intrinsic flaws considered should be based on the limits established under the manufacturing inspection and acceptance criteria and are expected to remain in service for the life of the structure.
(2) Impact Damage, which may occur during manufacturing and in service based on an evaluation of the threats by means of an impact survey and/or service experiences. This type of damage can include dents, penetrations, gouges, abrasions, and scratches. 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 Page C - 105
AC 27-1B, Chg 4 12/1/2011 of impact tests performed with configured structure, which is subjected to boundary conditions characteristic of the structure. Many different impact scenarios and locations are typically considered in the survey, which has a goal of identifying the most critical impacts (i.e., those causing the most serious damage but are least detectable). When simulating accidental impact damage, blunt or sharp impactors should be selected to represent the maximum criticality versus detectability, according to the load conditions
(e.g., tension, compression or shear). Until sufficient service experience exists to make good engineering judgments on energy and impactor variables, impact surveys should consider a wide range of conceivable impacts, including runway or ground debris, hail, tool drops, and vehicle collisions. Service data collected over time, can better define impact surveys and design criteria for subsequent products, as well as establish more rational inspection intervals and maintenance practice. Refer to paragraph f. (6)(ii)(H) for various combinations of detectability and energy levels to be considered in the damage tolerance and fatigue evaluation.
(3) Discrete Source Damage. The structure should be able to withstand limit static loads (considered as ultimate loads) and fatigue loads, which are reasonably expected during a completion of a flight on which damage resulting from obvious discrete source occurs (e.g., hail damage, bird strike, uncontained engine failure, and uncontained high energy rotating machinery failure). The extent of damage should be based on a rational assessment of service mission and potential damage relating to each discrete source.
(D) 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 (see AC 21-26). Assuming good process and quality controls, service history has shown that additional damage tolerant design considerations are also needed to ensure the safety of structure with secondary bonds
(i.e., random, but an unacceptable number 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 or a combination of the following:
(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 (l.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.
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(3) Critical bonded joints that have high static margins of safety (e.g., some rotor blades) may be accepted based on satisfactory service history of like or similar components.
(E) 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.
(F) Environmental effects (temperature and humidity representative of the expected service usage) on the static and 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.
(G) Variability in fatigue behavior should be covered by appropriate load or life scatter factors and these factors should take into account the number of specimens (H) The following Figure AC 27.573-2 illustrates the extent of the impact damage that needs to be considered in the damage tolerance and fatigue evaluation.
Detectability Obvious Damage Detectable damage Zone 2 Barely Detectable Impact Damage + Undetectable damage Zone 1 Zone 3 Energy Level of energy selected for complianec with §§ 27.305, 27.307 (Static Strength) and f(5)(vi) of this AC.
Maximum level of energy selected for the Damage Tolerance evaluation. This level should not be exceeded in serviec.
Figure AC 27.573-2. 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 Page C - 107