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.
(vili) 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.
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).
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
ELEMENTS SKIN
• WEB
GENERIC SPECIMENS COUPONS DATA BASE Figure AC 27 MG 8-1: Schematic diagram of building block tests.
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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.
(1) 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). I he 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 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 §§ 29.305, 307 (Static Strength) and g(5)(vi) of this AC paragraph Maximum level of energy selected for the Damage Tolerance evaluation This level should not be exceeded in serviec 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
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.
(6) Three Zones are defined by Figure AC 27 MG 8-2:
• Zone 1: Since the damage is not detectable, Ultimate Load capability is required.
The provisions of paragraph g(5) above provide a means of compliance.
• Zone 2: Since the damage can be detected at scheduled inspection, Limit Load (considered as Ultimate) capability is the minimum requirement for this damage.
• Zone 3: Since the damage is not detectable with the proposed in-service inspection procedures, ultimate load capability is required, unless an alternate procedure can show an equivalent level of safety. For example, residual strength lower than ultimate may be used in association with an improved inspection procedure.
(¡¡i) FATIGUE TOLERANCE EVALUATION - MEANS OF COMPLIANCE.
One, or a combination of, the methods below should show compliance with the requirements of this section. The Flaw Tolerant Safe-Life Evaluation or the Fail-Safe Evaluation are to be used unless it can be shown that neither can be achieved within the limitations of geometry, inspectability, or good design practice. In that case, the Safe-Life Evaluation should be used. From current state-of-the-art with rotorcraft applications, it is widely admitted that composite materials have good flaw or damage tolerance capabilities and therefore the safe-life option is rarely necessary. Flaw Tolerance evaluations are best suited for most composite structures, particularly those with structural redundancy and inherent resistance to damage growth. Damage resulting from anomalous or accidental events must be considered in the Flaw Tolerant Safe-Life and Fail-Safe evaluations.
The fatigue substantiation should include sufficient coupon, element, sub-element, or component tests to establish the fatigue scatter, curve shapes, and the environmental effects. The substantiation should include full-scale testing but also may be accomplished by analysis supported by test evidence. When spectrum testing is used, the lowest load levels can be eliminated from the spectrum if they can be shown to be non-damaging. The substantiation should include a static strength evaluation to show that the required residual strength and adequate stiffness, accounting for the effects of environment, are retained for the life of the structure or the appropriate inspection Page MG 8 - 21
AC 27-1B, Chg 2 4/25/06 interval. Flaws and damage as determined in paragraph g(6)(ii) above for the specific structure being substantiated should be imposed at each critical area of the structure.
(A) Flaw Tolerant Safe-Life Evaluation. This is a "No-Growth" method in that it demonstrates that the structure, with flaws present, is able to withstand repeated loads of variable magnitude without detectable flaw growth for the life of the rotorcraft or within a specified replacement time. This fatigue evaluation may be used to substantiate any type of damage that will remain in-service for the life of the structure.
No specific inspection requirements are generated from the test program in this method.
However, routine inspections for cracking, delaminations, and service damage as outlined in § 27.1529 are always required. Compliance using full-scale, component, or sub-component fatigue testing can be accomplished by either of the following methods:
(1) 'S-N Method. This method is based on determining the point where initiation of growth occurs for the flaws present at critical locations in the structure.
AC 27-1B, Chg 1, MG-11, provides guidance that can be appropriate for this method in composites. The method utilizes one or more full-scale, component, or sub-component test specimens subjected to constant-amplitude or spectrum loading applied in a distribution on the structure that is representative of critical flight conditions. Any indication of growth of the imposed flaws and defects, or structurally significant cracking, disbonding, splintering or delamination of the composite, defines the fatigue initiation characteristic of the structure in terms of applied load and cycles. Working S-N curves are established from the mean curve using strength or cycle reductions to account for fatigue scatter and environmental effects. Flight loads are compared to this working curve, and if any intercepts occur, a cumulative damage calculation is conducted to establish the component retirement time. Compliance with the ultimate load requirements should be demonstrated at the completion of the fatigue test.
(2) Life Test Method. This method uses spectrum fatigue testing to verify the absence of flaw growth over a large number of cycles that are equivalent to a lifetime of expected usage. The method uses one or more full-scale, component, or sub-component test specimens subjected to spectrum fatigue loading applied in a representative distribution of flight loads, including Ground-Air-Ground (GAG) loads.
Fatigue test loads should be increased by factors for environment and fatigue strength scatter. The load may also be increased using an S-N curve approach to reduce the duration of the test. Please reference "Certification Testing Methodology for Composite Structure, Report No. DOT/FAA/CT-86/39 for a discussion of the S-N approach. Any significant growth of the imposed flaws and defects, or structurally significant cracking, disbonding, splintering, or delamination of the composite during the test constitutes failure to achieve the desired lifetime. However, the equivalent life demonstrated at the time of inception of flaw growth or cracking can be used as a retirement time for the component. Compliance with the ultimate load requirements should be demonstrated at the completion of the fatigue test.
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(B) Fail-Safe (Residual Strength after Flaw Growth) Evaluation. This method demonstrates that the structure following a partial failure still has a sufficient residual strength capability within a specified inspection interval or the established retirement life of the component. If a retirement life is established, an ultimate design load capability is generally required while, if an inspection interval is determined, a limit load capability is the minimum acceptable residual strength capability that needs to be demonstrated. Full-scale, component, or sub-component testing should be accomplished using one or more specimens subjected to constant amplitude or spectrum loading applied in a manner representative of flight load conditions. The test loads should be increased by factors that account for environment and fatigue strength scatter. The results of the testing can be used to manage the structure in one of the three methods described below or a combination thereof.
(1) Fail-Safe, No Growth Evaluation. This approach is appropriate for inspectable in-service accidental damage. Structural details, elements, and subcomponents of critical structural areas, components, or full-scale structures, should be tested under repeated loads for validating a no-growth approach to the flaw tolerance requirements. The number of cycles applied to validate a no-growth concept should be statistically significant, and may be determined by load or life considerations. Residual strength testing or evaluation should be performed after repeated load cycling and demonstrate that the residual strength of the structure is equal to or greater than limit load considered as ultimate. Moreover, it should be shown that stiffness properties have not changed beyond acceptable levels. Inspection intervals should be established, considering the residual strength capability associated with the assumed damage. The intent of this is to assure that structure is not exposed to an excessive period of time with static margins less than ultimate, providing a lower safety level than in the typical slow growth situation, as illustrated in the Figure AC 27 MG 8-3. Once the damage is detected, the component is either repaired to restore ultimate load capability or replaced.
Residual strength Damage initiation or occurrence Slow-growth approach * Ultimate loads - No-growth approach ** Limit loads * Repair to Restore Ultimate Strength ** No growth without repair is not acceptable Time Figure AC 27 MG 8-3: Residual Strength vs. Time Page MG 8 - 23