AC 29-2C, Chg 4 12/1/2011
(viii) 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 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 § 29.609.
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 § 29.610.
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 § 29.610), 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 §§ 29.305 and
29.307 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.
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(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 (see paragraph f.(2)) and protection of structure (see paragraph f.(3)). Static strength tests 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 29.573-1 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
29.573-1, 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 29-2C, 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 29.573-1: Schematic Diagram of Building Block Tests.
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(iii) Allowables should be evaluated and used as specified in § 29.613.
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 (§ 29.619) 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 § 29.573.
(i) Background. The static strength determination required by §§ 29.305 and
29.307 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 § 29.573 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 failure. 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 Page C - 147
AC 29-2C, Chg 4 12/1/2011 manufacturing acceptance criteria. The damage tolerance assessment establishes 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 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.
All of the damages identified in the preceding paragraph (B) above should be derived from the threat assessment described in the following paragraph (C).
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(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 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, well-controlled, 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 resinstarved 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, goug es, abrasions, and scratches. A threat assessment is needed to identify impact damage Page C - 149
AC 29-2C, Chg 4 12/1/2011 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 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 (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).
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(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 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 tested.
(H) The following Figure AC 29.573-2 illustrates the extent of the impact damage that needs to be considered in the damage tolerance and fatigue evaluation.
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Detectability Obvious Damage ~ ~ Detectable damage [ Barely Detectable Impac~ Damage + Undetectable damage + I Zone 1 I + Level of energy selected for compliance with §§ 29.305, 29.307 (Static Strength) and fl:5)(vi) of this AC.
Maximum level of energy selected for the Damage Tolerance evaluation. This level should not be exceeded in service.
Energy AC 29-2C, Chg 4 12/1/2011 Figure AC 29.573-2: 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 (depicted in Figure AC 29.573-2) are dependent on the part of the structure under evaluation and a threat assessment.
(2) Obvious impact damage is used to define the threshold from which damage is readily detectable and appropriate actions may be taken before the next flight.
(3) Barely Detectable Impact Damage (BDID) is the state of damage at the threshold of detectability for the approved inspection procedure. Barely Visible Impact Damage (BVID) is that threshold of visually detectable damage associated with a detailed visual inspection procedure.
(4) Detectable Damage is the state of damage that can be reliably detected at scheduled inspection intervals. Visible Impact Damage (VID) is that threshold associated with the type of damage that should be detectable during a detailed visual inspection.
(5) Three Zones are depicted by this figure:
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Zone 1: Since the damage is not detectable, Ultimate Load capability is required. The provisions of paragraph f.(5) provide a means of compliance.
Zone 2: Since the damage can be detected at a scheduled inspection, Limit Load (considered as Ultimate load) capability is the minimum requirement for this damage.
Zone 3: Since the damage is not detectable with the proposed inservice 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 improved inspection procedures or with a probabilistic approach showing that the occurrence of energy levels is low enough so that an acceptable level of safety can be achieved.
Of the three zones, only Zone 3 may have a residual strength requirement that can vary with alternate procedures or the probability of damage occurrence or both. In either case, any compromise for residual strength requirements less than the ultimate load requirement should only be considered when pursuing one of the options under the damage tolerant fail-safe means of compliance, as described in the following section,
f.(6)(iii)(B). One example of the use of alternate procedures is for the rare damage threat from a high energy, blunt impact (e.g., service vehicle collision). Depending on the selected maintenance inspection scheme, such damage may fall under Zone 3.
When considering such damage in the design of a part, it may be shown to be damage tolerant fail safe, even though the damage is not detectable, based on a very low probability of occurrence. As a result, the design may have sufficiently high residual strength (e.g., below Ultimate, but well above limit load capability to ensure safety without detection for long periods of time). If it is further determined that such impact events usually occur with the knowledge of maintenance or aircraft service personnel, then the alternate procedures may be added to the Instructions for Continued Airworthiness. For example, advanced inspection methods, which can detect damage from high-energy blunt impacts, may be used as alternate procedures to minimize the risk of catastrophic failure for such Zone 3 damage.
(iii) Means of compliance. For each PSE, inspections, replacement times, or other procedures must be established as necessary to avoid catastrophic failure.
Compliance with the requirements of § 29.573(d) and (e) should be shown by one, or a combination of, the methods described subsequently. Generally, replacement times are established using Damage Tolerance Safe Life Evaluations and Inspection Intervals are established using Fail Safe Evaluations. From current state- of-the-art rotorcraft applications, it is widely accepted that composite materials have good flaw and damage tolerance capabilities and therefore the supplemental procedures may only be rarely necessary. Damage tolerance evaluations are best suited for 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 damage tolerant evaluations. The damage tolerant evaluation for replacement times and inspection intervals is to be used unless it is established that neither can be Page C - 153
AC 29-2C, Chg 4 12/1/2011 achieved within the limitations of geometry, inspectability, or good design practice. In that case, supplemental procedures must be established and submitted to the FAA for approval. In any case, the FAA must approve the methodology used for compliance to § 29.573. The substantiation method(s) should be chosen so that the structure is protected against catastrophic failure from each of the threats identified in paragraph
f.(6)(ii)(C) of this AC by a specific procedure (inspection, replacement time, or other procedure). For example, a manufacturing -related void of a specific allowable size could be substantiated by means of a replacement time method with no scheduled inspection. An accidental impact in the same area could be substantiated by an inspection method with no specific replacement time. The result could be one structure with several different inspection requirements (location, method, and interval) and a fixed replacement time as well. This combination of procedures assures that each threat is covered. 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, component, or subcomponent fatigue 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 interval. Damage as determined in paragraph f.(6)(ii) of this AC for the specific structure being substantiated should be imposed at each critical area of the structure.
(A) Damage Tolerant Safe-Life Evaluation. This is a “No -Growth” method in which the structure, with damage present, is able to withstand repeated loads of variable magnitude without detectable damage growth for the life of the rotorcraft or within a specified replacement time. This evaluation may be used to substantiate any type of damage that will remain in-service for the life of the part. No specific inspection requirements are generated from the test program in this method. However, compliance with routine inspections for cracking, delaminations, and service damage and other limitations prescribed in accordance with § 29.1529 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 damage present at critical locations in the structure.
AC 27-1B, AC 27 MG 11, provides guidance that may be appropriate for this method.
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 damage and defects, or structurally significant cracking, disbonding, splintering, or delaminating 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 or both to account for fatigue scatter and environmental effects. Flight loads are compared to this Page C - 154
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 damage 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. Any significant growth of the imposed damage, 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 damage 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.
(B) Damage Tolerant Fail-Safe (Residual Strength with Detectable Damage) Evaluation. This method establishes inspection intervals to ensure that the structure remaining after a partial failure is able to withstand design limit loads without failure or excessive structural deformations within a specified inspection interval. If the damage is detected in an inspection, the structure should be either replaced or repaired to restore ultimate load capability. Evaluation of Zone 3 damage should have sufficiently high residual strength and, if necessary, supplemental procedures should be established to minimize the risk of catastrophic failure. Full- scale, component, or subcomponent 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 or a combination of the three methods described subsequently.
(1) No Growth Evaluation. This approach is appropriate for inspectable in-service damage, which does not grow in service (see Figure AC 29.573-3). Damage growth should be substantiated using either method described in f.(6)(iii)(B)(2) or
f.(6)(iii)(B)(3). Structural details, elements, sub-comp onents, and components of critical structural areas, or full-scale structures, should be tested under repeated loads for validating a no-growth approach to the damage 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 or both. Residual strength testing or evaluations should be performed after repeated load cycling demonstrating 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 Page C - 155