AC 20-107B

Advisory U.S. Department of Transportation Circular Federal Aviation Administration Subject: Change 1 to COMPOSITE Date: August 24, 2010 AC No: 20-107B AIRCRAFT STRUCTURE Initiated by: AIR-100 Change: 1

1. PURPOSE. This Change corrects minor errors in the original AC. Under paragraph 8a(1)(c)(i), an "or" is changed to "and"; under Figure 3 in paragraph 8a(1)(c), the box for "Category 3 Damage" is reworded; the definition for "No-Growth Approach" is moved to place it into correct alphabetical order; and page numbers within Appendix 2 are revised to be correct.

2. CHANGE TEXT. Changed text is indicated by vertical bars in the margins.

PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated 12 9/8/09 12 8/24/10 13 9/8/09 13 8/24/10 Appendix 2, A2-2 9/8/09 Appendix 2, A2-2 8/24/10 Appendix 2, A2-3 19/8/09 Appendix 2, A2-3 8/24/10 D091e David W. Hempe Manager, Aircraft Engineering Division Aircraft Certification Service

Advisory U.S. Department of Transportation CircularFederal Aviation Administration Subject: COMPOSITE AIRCRAFT Date: 9/8/09 AC No: 20-107B STRUCTURE Initiated by: AIR-100 Change:

1. Purpose of this Advisory Circular (AC). This AC sets forth an acceptable means, but not the only means of showing compliance with the provisions of Title 14 of the Code of Federal Regulations (14 CFR) parts 23, 25, 27, and 29 regarding airworthiness type certification requirements for composite aircraft structures involving fiber reinforced materials, e.g., carbon and glass fiber reinforced plastics. Guidance information is also presented on the closely related design, manufacturing, and maintenance aspects. The information contained herein is for guidance purposes and is not mandatory or regulatory in nature.

2. To Whom this AC Applies. The audience of this AC may include applicants, certificate/approval holders, parts manufacturers, material suppliers, maintenance, and repair organizations.

3. Cancellation. AC 20-107A, Composite Aircraft Structure, dated April 25, 1984, is cancelled.

4. Related Regulations and Guidance. The material contained herein applies to normal, utility, acrobatic, commuter, and transport category aircraft type certificated under Civil Aviation Regulations (CARs) 3, 4b, 6, 7; and 14 CFR parts 23, 25, 27, 29; and it is produced in compliance with 14 CFR part 21, §§ 21.125, or 21.143 as may be appropriate. The sections of 14 CFR, parts 23, 25, 27, and 29 applicable to each paragraph of this AC are listed in Appendix

1. Other supporting guidance relevant to the AC is also provided in Appendix 1.

5. General.

a. The procedures outlined in this AC provide guidance material for composite structures, particularly those that are essential in maintaining the overall flight safety of the aircraft (“critical structure” as defined in Appendix 2), and are considered acceptable to the FAA for showing compliance with certification requirements of civil composite aircraft. This circular is published to aid in the evaluation of certification programs for composite applications and to reflect the current status of composite technology. It is expected that this circular will be modified

periodically to reflect the continued evolution of composite technology and the data collected from service experience and expanding applications.

b. There are factors unique to the specific composite materials and processes used for a given application. For example, the environm ental sensitivity, anisotropic properties, and heterogeneous nature of composites can make the determination of structural failure loads, modes, and locations difficult. The reliabili ty of such evaluation depends on repeatable structural details created by scaled manufacturing or repair processes. The extent of testing and/or analysis may differ for a structure depending upon the criticality to flight safety, expected service usage, the material and processes selected, the design margins, the failure criteria, the database and experience with similar structures, and on other factors affecting a particular structure. It is expect ed that these factors will be considered when interpreting this AC for use on a specific application.

c. Definitions of terms used in this AC can be found in Appendix 2.

6. Material and Fabrication Development. All composite materials and processes used in structures are qualified through enough fabrication trials and tests to demonstrate a reproducible and reliable design. One of the unique features of composite construction is the degree of care needed in the procurement and processing of composite materials. The final mechanical behavior of a given composite material may vary greatly depending on the processing methods employed to fabricate production parts. Special care needs to be taken in controlling both the materials being procured and how the material is processed once delivered to the fabrication facility. Regulatory requirements in 14 CFR, parts 2X, §§ 2x.603 and 2x.605 specify the need to procure and process materials under approved material and process specifications that control the key parameters governing performance. 14 CFR, parts 2X, §§ 2x.609 and 2x.613 outlines a need to protect structures against the degradation possible in service. They also require that the design account for any changes in performance (e.g., environmental and variability effects) permitted by material and process specifications.

a. Material and Process Control.

(1) Specifications covering material, material processing, and fabrication procedures are established to ensure a basis for fabricating reproducible and reliable structure. Material specifications are required to ensure consistent material is being procured, and batch acceptance testing or statistical process controls are used to ensure material properties do not drift over time.

Specifications covering processing procedures should be developed to ensure that repeatable and reliable structure is being manufactured. The means of processing qualification and acceptance tests defined in each material specification should be representative of the expected applicable manufacturing process. The process parameters for fabricating test specimens should match the process parameters used in manufacturing actual production parts as closely as possible. Both test and production parts must conform to material and process specifications.

(2) Once the fabrication processes have been established, changes should not occur unless additional qualification, including testing of differences is completed (refer to Appendix Page 2

3). It is important to estab lish processing tolerances; material handling and storage limits; and key characteristics, which can be measured and tracked to judge part quality.

(3) Material requirements identified in procurement specifications should be based on the qualification test results for samples produced using the related process specifications.

Qualification data must cover all properties important to the control of materials (composites and adhesives) and processes used for production of composite structure. Carefully selected physical, chemical, and mechanical qualification tests are used to demonstrate the formulation, stiffness, strength, durability, and reliability of materials and processes for aircraft applications.

It is recommended that material suppliers work closely with airframe manufacturers to properly define material requirements.

(4) To provide an adequate design database, environmental effects on critical properties of the material systems and associated processes should be established. In addition to testing in an ambient environment, variables should include extreme service temperature and moisture content conditions and effects of long-term durability. Qualification tests for environmental effects and long-term durability are particularly important when evaluating the materials, processes, and interface issues associated with structural bonding (refer to paragraph 6.c for related guidance).

(5) Key characteristics and processing parameters will be monitored for in-process quality control. The overall quality control plan required by the certifying agency should involve all relevant disciplines, i.e., engineering, manufacturing, and quality control. A reliable quality control system should be in place to address special engineering requirements that arise in individual parts or areas as a result of potential failure modes, damage tolerance and flaw growth requirements, loadings, inspectability, and local sensitivities to manufacture and assembly.

(6) The discrepancies permitted by the specifications should also be substantiated by analysis supported by test evidence, or tests at the coupon, element or subcomponent level. For new production methods, repeatable processes should be demonstrated at sufficient structural scale in a way shown to be consistent with the material and process qualification tests and development of the associated specifications. This will require integration of the technical issues associated with product design and manufacturing details prior to a large investment in structural tests and analysis correlation. It will also ensure the relevance of quality control procedures defined to control materials and processes as related to the product structural details.

(7) The FAA does not generally certify materials and processes. However, the materials and processes may be accepted as part of a particular aircraft product certification. Appropriate credit may be given to organizations using the same materials and processes in similar applications subject to substantiation and applicability. In some cases, material and processing information may become part of accepted shared databases used throughout the industry. New users of shared qualification databases must control the associated materials and processes through proper use of the related specifications and demonstrate their understanding by performing equivalency sampling tests for key properties. Materials and processes used in technical standard order (TSO) articles or authorizations must also be qualified and controlled.

b. Manufacturing Implementation.

(1) Process specifications and manufacturing documentation are needed to control composite fabrication and assembly. The environment and cleanliness of facilities are controlled to a level validated by qualification and proof of structure testing. Raw and ancillary materials are controlled to specification requirements that are consistent with material and process qualifications. Parts fabricated meet the producti on tolerances validated in qualification, design data development, and proof of structure tests. Some key fabrication process considerations requiring such control include material handling and storage; laminate layup and bagging (or other alternate process steps for non-laminated material forms and advanced processes); mating part dimensional tolerance control; part cure (thermal management); machining and assembly;

cured part inspection and handling procedures; and technician training for specific material, processes, tooling and equipment.

(2) Thorough manufacturing records are needed to support parts acceptance and allowable discrepancies (defects, damage and anomalies). Substantiating data is needed to justify all known defects, damage and anomalies allowed to remain in service without rework or repair. Manufacturing records are also needed for all substantiated design and process changes.

(3) New suppliers of parts for previously certified aircraft products are qualified by manufacturing trials and quality assessments to ensure equivalent production and repeatability.

Some destructive inspection of critical structural details is needed for manufacturing flaws that are not end item inspectable and require process controls to ensure reliable fabrication.

c. Structural Bonding. Bonded structures include multiple interfaces (e.g., composite-tocomposite, composite-to-metal, or metal-to-metal), where at least one of the interfaces requires additional surface preparation prior to bonding. The general nature of technical parameters that govern different types of bonded structures are similar. A qualified bonding process is documented after demonstrating repeatable and reliable processing steps such as surface preparation. It entails unders tanding the sensitivity of structural performance based upon expected variation permitted per the process. Characterization outside the process limits is recommended to ensure process robustness. In the case of bonding composite interfaces, a qualified surface preparation of all previously cured substrates is needed to activate their surface for chemical adhesion. All metal interfaces in a bonded structure also have chemically activated surfaces created by a qualified preparation process. Many technical issues for bonding require cross-functional teams for successful applications. Applications require stringent process control and a thorough substantiation of structural integrity.

(1) Many bond failures and problems in service have been traced to invalid qualifications or insufficient quality control of production processes. Physical and chemical tests may be used to control surface preparation, adhesive mixing, viscosity, and cure properties (e.g., density, degree of cure, glass transition temperature). Lap shear stiffness and strength are common mechanical tests for adhesive and bond process qualification. Shear tests do not provide a reliable measure of long-term durability and environmental degradation associated with poor bonding processes (i.e., lack of adhesion). Some type of peel test has proven more reliable for evaluating proper adhesion. Without chemical bonding, the so-called condition of a “weak Page 4

bond” exists when the bonded joint is either loaded by peel forces or exposed to the environment over a long period of time, or both. Adhesion failures, which indicate the lack of chemical bonding between substrate and adhesive materials, are considered an unacceptable failure mode in all test types. Material or bond process probl ems that lead to adhesion failures are solved before proceeding with qualification tests.

(2) Process specifications are needed to control adhesive bonding in manufacturing and repair. A “process control mentality,” which includes a combination of in-process inspections and tests, has proven to be the most reliable means of ensuring the quality of adhesive bonds.

The environment and cleanliness of facilities used for bonding processes are controlled to a level validated by qualification and proof of structure testing. Adhesives and substrate materials are controlled to specification requirements that are consistent with material and bond process qualifications. The bonding processes used for pr oduction and repair meet tolerances validated in qualification, design data development, and proof of structure tests. Some key bond fabrication process considerations requiring such control include material handling and storage;

bond surface preparation; mating part dimensional tolerance control; adhesive application and clamp-up pressure; bond line thickness control; bonded part cure (thermal management); cured part inspection and handling procedures; and bond technician training for specific material, processes, tooling, and equipment. Bond surface preparation and subsequent handling controls leading up to the bond assembly and cure must be closely controlled in time and exposure to environment and contamination.

(3) 14 CFR § 23.573(a) sets forth requirements for substantiating the primary composite airframe structures, including considerations for damage tolerance, fatigue, and bonded joints.

Although this is a small airplane rule, the same performance standards are normally expected with transport and rotorcraft category aircraft (via special conditions and issue papers).

(a) For any bonded joint, § 23.573(a)(5) states in part: "the failure of which would result in catastrophic loss of the airplane, the limit load capacity must be substantiated by one of the following methods—(i) The maximum disbonds of each bonded joint consistent with the capability to withstand the loads in paragraph (a)(3) of this section must be determined by analysis, tests, or both. Disbonds of each bonded joint greater than this must be prevented by design features; or (ii) Proof testing must be conducted on each production article that will apply the critical limit design load to each critical bonded joint; or (iii) Repeatable and reliable non-destructive inspection techniques must be established that ensure the strength of each joint."

(b) These options do not supersede the need for a qualified bonding process and rigorous quality controls for bonded structures. For example, fail safety implied by the first option is not intended to provide adequate safety for the systematic problem of a bad bonding process applied to a fleet of aircraft structures. Instead, it gives fail safety against bonding problems that may occasionally occur over local areas (e.g., insufficient local bond contact pressure or contamination). Performing static proof tests to limit load, which is the second option, may not detect weak bonds requiring environmental exposure and time to degrade bonded joint strength. This issue should be c overed by adequately demonstrating that qualified bonding materials and processes have long-term environmental durability. Finally, the third option is open for future advancement and validation of non-destructive inspection (NDI) Page 5

technology to detect weak bonds, which degrade over time and lead to adhesion failures. Such technology has not been reliably demonstrated at a production scale to date.

(4) Adhesion failures found in production require immediate action to identify the specific cause and isolate all affected parts and assemblies for disposition. Adhesion failures discovered in service require immediate action to determine the cause, to isolate the affected aircraft, and to conduct directed inspection and repair. Depending on the suspected severity of the bonding problem, immediate action may be required to restore the affected aircraft to an airworthy condition.

d. Environmental Considerations. Environmental design criteria should be developed that identify the critical environmental exposures, including humidity and temperature, to which the material in the application under evaluation may be exposed. Service data (e.g., moisture content as a function of time in service) can be used to ensure such criteria are realistic. In addition, the peak temperatures for composite structure installed in close proximity to aircraft systems that generate thermal energy need to be identified for worst-case normal operation and system failure cases. Environmental design crite ria are not required where existing data demonstrate that no significant environmental effects, including the effects of temperature and moisture, exist for the material system and construction details, within the bounds of environmental exposure being considered.

(1) Experimental evidence should be provided to demonstrate that the material design values or allowables are attained with a high degree of confidence in the appropriate critical environmental exposures to be expected in service. It should be realized that the worst case environment may not be the same for all structural details (e.g., hot wet conditions can be critical for some failure modes, while cold dry conditions may be worse for others). The effect of the service environment on static strength, fatigue and stiffness properties and design values should be determined for the material system through tests; e.g., accelerated environmental tests, or from applicable service data. The maximum moisture content considered is related to that possible during the service life, which may be a function of a given part thickness, moisture diffusion properties and realistic environmental exposures. The effects of environmental cycling

(i.e., moisture and temperature) should be evaluated when the application involves fluctuations or unique design details not covered in the past. Existing test data may be used where it can be shown to be directly applicable to the material system, design details, and environmental cycling conditions characteristic of the application. All accelerated test methods should be representative of real-time environmental and load exposure. Any factors used for acceleration that chemically alter the material (e.g., high temperatures that cause post-cure) should be avoided to ensure behavior representative of real environmental exposures.

(2) Depending on the design configuration, local structural details, and selected processes, the effects of residual stresses that depend on environment must be addressed (e.g., differential thermal expansion of attached parts).

e. Protection of Structure. Weathering, abrasion, erosion, ultraviolet radiation, and chemical environment (glycol, hydraulic fluid, fuel, cleaning agents, etc.) may cause deterioration in a composite structure. Suitable protection against and/or consideration of Page 6

degradation in material properties should be provided for conditions expected in service and demonstrated by test. Where necessary, provide provisions for ventilation and drainage.

Isolation layers are needed at the interfaces between some composite and metal materials to avoid corrosion (e.g., glass plies may be used to isolate carbon composite layers from aluminum). In addition, qualification of the special fasteners and installation procedures used for parts made from composite materials need to address the galvanic corrosion issues, as well as the potential for damaging the composite (delamination and fiber breakage) in forming the fastener.

f. Design Values. Data used to derive design values must be obtained from stable and repeatable material, which are procured per a mature material specification and processed per a representative production process specification. This is done to ensure that the permitted variability of the production materials is captured in the statistical analysis used to derive the design values. Design values derived too early in the material’s development stage, before raw material and composite part production processes have matured, may not satisfy the intent of the associated rules. Laminated material system de sign values should be established on the laminate level by either test of the laminate or by test of the lamina in conjunction with a test validated analytical method. Similarly, design values for non-laminated material forms and advanced composite processes must be established at the scale that best represents the material as it appears in the part or by tests of material substructure in conjunction with a test validated analytical method.

g. Structural Details. For a specific structural configuration of an individual component (point design), design values may be established which include the effects of appropriate design features (holes, joints, etc.). Specific metrics that quantify the severity of composite structural damage states caused by foreign impact damage threats are needed to perform analysis (i.e., the equivalent of a metallic crack length). As a result, testing will often be needed to characterize residual strength, including the structural effects of critical damage location and combined loads.

Different levels of impact damage are generally accommodated by limiting the design strain levels for ultimate and limit combined load design criteria. In this manner, rational analyses supported by tests can be established to characterize residual strength for point design details.

7. Proof of Structure – Static. The structural static strength substantiation of a composite design should consider all critical load cases and associated failure modes. It should also include effects of environment (including structural residual stresses induced during the fabrication process), material and process variability, non-detectable defects or any defects that are allowed by the quality control, manufacturing acceptance criteria, and service damage allowed in maintenance documents of the end product. The static strength of the composite design should be demonstrated through a program of component ultimate load tests in the appropriate environment, unless experience with similar designs, material systems, and loadings is available to demonstrate the adequacy of the analysis supported by subcomponent, element and coupon 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.

a. The effects of repeated loading and environmental exposure which may result in material property degradation should be addressed in the static strength evaluation. This can be shown by Page 7

analysis supported by test evidence, by tests at the coupon, element or subcomponent level, as appropriate, or alternatively by relevant existing data. Earlier discussions in this AC address the effects of environment on material properties (paragraph 6.d) and protection of structure (paragraph 6.e). For critical loading conditions, three approaches exist to account for prior repeated loading and/or environmental exposure in the full scale static test.

(1) In the first approach, the full scale static test should be conducted on structure with prior repeated loading and conditioned to simulate the critical environmental exposure and then tested in that environment.

(2) The second approach relies upon coupon, element, and subcomponent test data to determine the effect of repeated loading and environmental exposure on static strength. The degradation characterized by these tests should then be accounted for in the full scale static strength demonstration test (e.g., overload factors), or in analysis of these results (e.g., showing a positive margin of safety with design values that include the degrading effects of environment and repeated load).

(3) In practice, aspects of the first two approaches may be combined to get the desired result (e.g., a full scale static test may be performed at critical operating temperature with a load factor to account for moisture absorbed over the aircraft structure’s life). Alternate means to account for environment using validated tests and analyses (e.g., an equivalent temperature enhancement to account for the effect of moisture without chemically altering the material) may be proposed by the applicant to the administrator for approval.

b. The strength of the composite structure should be reliably established, incrementally, through a program of analysis and a series of tests conducted using specimens of varying levels of complexity. Often referred to in industry as the “building block” approach, these tests and analyses at the coupon, element, details, and 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. Typically, testing progresses from simple specimens to more complex elements and details over time. This approach allows the data collected for sufficient analysis correlation and the necessary replicates to quantify variations occurring at the larger structural scales to be economically obtained. The lessons learned from initial tests also help avoid early failures in more complex full scale tests, which are more costly to conduct and often occur later in a certification program schedule.

(1) Figures 1 and 2 provide a conceptual schematic of tests typically included in the building block approach for fixed wing and rotor blade structures, respectively. 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 subcomponent and component tests. Detail and subcomponent tests may be used to validate the ability of analysis methods to predict local strains and failure modes. Additional statistical considerations (e.g., repetitive point design testing and/or component overload factors to cover material and process variability) will be needed when analysis validation is not achieved. The static strength substantiation program should also consider all critical loading conditions for all critical structure. This includes an assessment of residual strength and stiffness requirements Page 8

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9/8/09 AC 20-107B after a predetermined length of service, which takes into account damage and other degradation due to the service period.

Figure 1 - Schematic diagram of building block tests for a fixed wing Figure 2 - Schematic diagram of building block tests for a tail rotor blade

(2) Successful static strength substantiation of composite structures has traditionally depended on proper consideration of stress concentrations (e.g., notch sensitivity of details and impact damage), competing failure modes, and out-of-plane loads. A complete building block approach to composite structural substantiation addresses most critical structural issues in test articles with increasing levels of complexity such that many areas of reliable performance can be demonstrated prior to the component tests. The details and subcomponent testing should establish failure criteria and account for impact damage in assembled composite structures.

Component tests are needed to provide the final validation accounting for combined loads and complex load paths, which include some out-of-plane effects. When using the building block approach, the critical load cases and associated failure modes would be identified for component tests using the analytical methods, which are supported by test validation.

c. The component static test may be performed in an ambient atmosphere if the effects of the environment are reliably predicted by building block tests and are accounted for in the static test or in the analysis of the results of the static test.

d. The static test articles should be fabricated and assembled in accordance with production specifications and processes so that the test articles are representative of production structure including defects consistent with the limits established by manufacturing acceptance criteria.

e. The material and processing variability of the composite structure should be considered in the static strength substantiation. This is primarily achieved by establishing sufficient process and quality controls to manufacture structure and reliably substantiate the required strength by test and analysis. The scatter in strength properties due to variability in materials and processes are characterized by proper allowables or design values, which are derived in compliance with 14 CFR § 2x.613. When the detail, subcomponent and component tests show that local strains are adequately predicted, and positive margins of safety exist using a validated analysis everywhere on the structure, then proof of static strength is said to be substantiated using analysis supported by test evidence. Alternatively, in the absence of sufficient building block test data and analysis validation, overloads are needed in the component test to gain proof of static strength for the structure using an approach referred to as substantiated by tests. The overload factors applied in this case need to be substantiated either through tests or past experience and must account for the expected material and process variation.

f. It should be shown that impact damage that can be likely 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 can be shown by analysis supported by test evidence, or by a combination of tests at the coupon, element, subcomponent, and component levels. The realistic test assessment of impact damage requires proper consideration of the structural details and boundary conditions. When using a visual inspection procedure, the likely impact damage at the threshold of reliable detection has been called barely visible impact damage (BVID). Selection of impact sites for static strength substantiation should consider the criticality of the local structural detail, and the ability to inspect a location. The size and sh ape of impactors used for static strength substantiation should be consistent with likely impact damage scenarios that may go undetected for the life of an aircraft. Note that it is possible for some designs to have detectable impact damage and still meet static strength loads and other requirements without repair (refer to allowable damage discussions in paragraph 10.c.(1)).

g. Major material and process changes on existing certified structure require additional static strength substantiation (e.g., refer to Appendix 3).

8. Proof of Structure – Fatigue and Damage Tolerance. The evaluation of composite structure should be based on the applicable requirements of 14 CFR §§ 23.573(a), 25.571,

27.571, and 29.571. Such evaluation must show that catastrophic failure due to fatigue, environmental effects, manufacturing defects, or accidental damage will be avoided throughout the operational life of the aircraft. The nature and extent of analysis or tests on complete structures and/or portions of the primary structure will depend upon applicable previous fatigue/damage tolerant designs, construction, tests, and service experience on similar structures.

In the absence of experience with similar designs, FAA-approved structural development tests of components, subcomponents, and elements should be performed (following the same principles discussed in paragraph 7.b and Appendix 3). The following considerations are unique to the use of composite material systems and provide guidance for the method of substantiation selected by the applicant. When establishing details fo r the damage tolerance and fatigue evaluation, attention should be given to a thorough damage threat assessment, geometry, inspectability, good design practice, and the types of damage/degradation of the structure under consideration.

 Composite damage tolerance and fatigue performance is strongly dependent on structural design details (e.g., skin laminate stacking sequence, stringer or frame spacing, stiffening element attachment details, damage arrestment features, and structural redundancy).

 Composite damage tolerance and fatigue evaluations require substantiation in component tests unless experience with similar designs, material systems, and loadings is available to demonstrate the adequacy of the analysis supported by coupons, elements, and subcomponent tests.

 Final static strength, fatigue, and damage tolerance substantiation may be gained in testing a single component test article if sufficient building block test evidence exists to ensure that the selected sequence of repeated and static loading yield results representative of service or provide a conservative evaluation.

 Peak repeated loads are needed to practically demonstrate the fatigue and damage tolerance of composite aircraft structure in a limited number of component tests. As a result, metal structures present in the test article generally require additional consideration and testing. The info rmation contained in AC 25.571-1 provides fatigue and damage tolerance guidance for metallic structures.

a. Damage Tolerance Evaluation.

(1) Damage tolerance evaluation starts with identification of structure whose failure would reduce the structural integrity of the aircraft. A damage threat assessment must be performed for the structure to determine possible locations, types, and sizes of damage considering fatigue, environmental effects, intrinsic flaws, and foreign object impact or other accidental damage (including discrete source) that may occur during manufacture, operation or maintenance.