0 U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: Damage Tolerance Date: 1/13/2011 AC No: 25.571-1D and Fatigue Evaluation of Structure Initiated By: ANM-115
1. Purpose.
a. This advisory circular (AC) provides guidance for compliance with the provisions of Title 14, Code of Federal Regulations (14 CFR) part 25, pertaining to the requirements for damage-tolerance and fatigue evaluation of transport category aircraft structure, including evaluation of widespread fatigue damage (WFD) and establishing a limit of validity of the engineering data that supports the structural-maintenance program (hereafter referred to as the LOV). This AC also includes guidance pertaining to discrete source damage.
b. The following appendices appear at the end of this AC:
● Appendix 1 – References and Definitions ● Appendix 2 – Full-Scale Fatigue Testing ● Appendix 3 – Process for Compliance with 14 CFR 25.571(a), (b) and (c) for Fatigue Damage ● Appendix 4 – Examples of Alterations that May Require Full-Scale Fatigue Testing ● Appendix 5 – PSE, FCS and WFD-Susceptible Structure ● Appendix 6 – Acronyms and Abbreviations Used in this AC
2. Applicability.
a. The guidance provided in this document is directed to airplane manufacturers, modifiers, foreign civil-aviation authorities, and Federal Aviation Administration (FAA) transport airplane type certification engineers and designees.
b. This material is neither mandatory nor regulatory in nature and does not constitute a regulation. It describes acceptable means, but not the only means, for demonstrating compliance with the applicable regulations. The FAA will consider other methods of demonstrating compliance that an applicant may elect to present. While these guidelines are not mandatory, they are derived from extensive FAA and industry experience in determining compliance with the relevant regulations. On the other hand, if the FAA becomes aware of circumstances that convince us that following this AC would not result in compliance with the applicable 1
regulations, we will not be bound by the terms of this AC, and we may require additional substantiation or design changes as a basis for finding compliance.
c. This material does not change, create additional, authorize changes in, or permit deviation from regulatory requirements.
3. Cancellation. AC 25.571-1C, dated April 29, 1998, is canceled.
4. Background.
a. Since the early 1970s, there have been significant developments in the state-of-theart and industry-practice in the area of structural-fatigue and fail-safe strength evaluation of transport category airplanes. Recognizing that these developments could warrant some revision of the existing fatigue requirements of §§ 25.571 and 25.573, on November 18, 1976 (41 FR 50956), the FAA gave notice of the Transport Category Airplane Fatigue Regulatory Review Program, inviting interested persons to submit proposals to amend those requirements.
The proposals and related discussions formed the basis for the revision of the structural-fatigueevaluation standards of §§ 25.571 and 25.573, and the development of guidance material. To that end, § 25.571 was revised, § 25.573 was deleted (the scope of § 25.571 was expanded to cover the substance of the deleted section), and guidance material (AC 25.571-1) was prepared.
b. After issuance of AC 25.571-1 on September 28, 1978, additional guidance material, including guidance on discrete source damage, was developed and incorporated in revision 1A, issued on March 5, 1986. The AC was further re vised on February 18, 1997; revision 1B added guidance on the elements to be considered in developing scatter factors for certification.
Revision 1C of the AC was issued April 29, 1998, to add guidance pertaining to precluding widespread fatigue damage within the design service goal of the airplane and determining thresholds for fatigue inspection. This revisi on removes guidance material for precluding WFD up to the design service goal and replaces it with guidance material for precluding WFD up to the LOV.
5. Introduction.
a. General. The FAA considers the contents of this AC in determining compliance with the damage-tolerance, fatigue, and discrete source damage requirements of § 25.571.
(1) The requirements of § 25.571 apply equally to metallic and composite structure. The focus of this AC is metallic structure. Refer to AC 20-107B for guidance on composite structure.
(2) Section 25.571 requires applicants to evaluate all structure that could contribute to catastrophic failure of the airplane with respect to its susceptibility to fatigue, corrosion, and accidental damage. The applicant must establish inspections or other procedures (herein also referred to as maintenance actions) as necessary to avoid catastrophic failure during the operational life of the airplane based on the results of these evaluations. Section 25.571 also 2
requires the applicant to establish an LOV. The LOV, in effect, is the operational life of the airplane consistent with evaluations accomplished and maintenance actions established to prevent WFD. Although the LOV is established based on WFD considerations, it is intended that all maintenance actions required to address fatigue, corrosion, and accidental damage up to the LOV are identified in the structural-maintenance program. All inspections and other procedures (e.g., modification times, replacement times) that are necessary to prevent a catastrophic failure due to fatigue, up to the LOV, must be included in the Airworthiness Limitations section (ALS) of the Instructions for Continued Airworthiness (ICA), as required by § 25.1529, along with the LOV.
(3) Although a uniform approach to the evaluation required by § 25.571 is desirable, the FAA recognizes that, in such a complex field, new design features and methods of fabrication, new approaches to the evaluation, and new configurations could necessitate variations and deviations from the procedures described in this AC.
(4) Compliance with § 25.571(b), Damage-tolerance evaluation, is not required if the applicant establishes that it is impracticable to apply the regulation to particular structure.
An example of structure that may not be conducive to damage-tolerant design is the landing gear and its attachments. In this case, the applicant may comply instead with § 25.571(c), Fatigue (safe-life) evaluation. The FAA does not allow exclusive reliance on damage-tolerance-based inspections for structure susceptible to WFD. The applicant must demonstrate that WFD is unlikely to occur prior to the LOV for the airplane. An applicant may use damage-tolerancebased inspections for multiple site damage (MSD)/multiple element damage (MED) to supplement replacement or modification required to preclude WFD when shown to be practical.
(5) The applicant should perform crack-growth and residual-strength testing to produce the design data needed to support crack-growth and residual-strength analyses. Fullscale fatigue-test evidence is required to support the evaluation of structure that is susceptible to WFD. Test evidence is needed to support analysis used to establish safe-life replacement times.
(6) Replacement times, inspections, or other procedures to address fatigue cracking must be established as necessary. These actions must be based on quantitative evaluations of the fatigue characteristics of the structure. In general, analysis supported by test evidence will be necessary to generate the information needed; service experience may also be used. The applicant should identify the intended approach in a compliance plan. All inspections, modification times, replacement times, and LOVs that are necessary to prevent a catastrophic failure – based on the damage-tolerance, fatigue, and WFD evaluations – must be included in the ALS of the ICA as required by § 25.1529.
(7) The applicant must establish inspections or other procedures for environmental damage and accidental damage as necessary to prevent catastrophic failure.
Guidance for establishing environmental damage and accidental damage maintenance actions is included in section 6 of this AC.
b. Typical loading spectrum expected in service. The loading spectrum used for crack-growth and fatigue-crack-initiation assessments (tests or analyses) should be based on measured statistical data of the type derived from government and industry load-history studies 3
and, where data is insufficient or unavailable, on a conservative estimate of the anticipated use of the airplane. The principal loads that should be considered in establishing a loading spectrum are flight loads (gust and maneuver), ground loads (taxiing, landing impact, turning, engine run-up, braking, thrust reversing, and towing), and pressurization loads. The development of the loading spectrum includes the definition of the expected flight plan, which involves climb, cruise, descent, flight times, operational speeds and altitudes, and the approximate time to be spent in each of the operating regimes. Operations for crew training and other pertinent factors, such as the dynamic stress characteristics of any flexible structure excited by turbulence or buffeting, should also be considered. For pressurized cabins, the loading spectrum should include the repeated application of the normal operating differential pressure, and the superimposed effects of flight loads and external aerodynamic pressures.
c. Structure to be evaluated. When assessing the possibility of fatigue failures, the applicant should examine the design to determine probable points of failure in service. In this examination, consideration should be given, as necessary, to (1) the results of stress analyses, static tests, fatigue tests, strain-gauge surveys, and tests of similar structural configurations and
(2) service experience. Service experience has shown that special attention should be focused on the design details of important discontinuities, attachment fittings, tension joints, splices, and cutouts such as windows, doors, and other openings. Locations prone to accidental damage (such as that due to impact with ground servicing equipment near airplane doors) or to corrosion should also be considered.
d. Analyses and tests. Repeated load analyses or tests should be conducted on structures representative of components or subcomponents of the wing, control surfaces, empennage, fuselage, landing gear, and their related primary attachments. Analyses and tests need not be conducted if it is determined from the foregoing examination that the normal operating stresses in specific regions of the structure are of such a low order that significant damage growth is extremely improbable. Test articles should include structure representative of attachment fittings, tension joints, splices, changes in section, cutouts, and discontinuities. Any method used in the analyses should be supported, as necessary, by test or service experience.
Typical (average) values of material properties that account for the quantifiable effects of environment and other parameters may be used in residual-strength, crack-growth, and damagedetection analyses for damage-tolerance assessments and for discrete source damage. These are described, respectively, in sections 6 and 9 of this AC.
6. Damage-Tolerance Evaluation.
a. General. The damage-tolerance evaluation of structure is intended to ensure that— should fatigue, corrosion, or accidental damage occur within the LOV of the airplane—the remaining structure can withstand reasonable loads without failure or excessive structural deformation until the damage is detected. The damage-tolerance evaluation should include the following:
identification of the structure to be evaluated, definition of the loading conditions and extent of damage, 4
structural tests or analyses, or both, to substantiate that the design objective has been achieved, and generation of supporting data for inspection programs to ensure detection of damage.
Although this evaluation applies to either single- or multiple-load-path structure, the use of multiple load path structure should be given high priority in achieving a damage-tolerant design. Design features that s hould be considered in attaining a damage-tolerant structure include the following:
(1) Multiple load path construction, and the use of crack stoppers to control the rate of crack growth and to provide adequate residual strength;
(2) Materials and stress levels that, after initiation of cracks provide a controlled slow rate of crack propagation combined with high residual strength;
(3) Arrangement of design details to ensure a high probability that a failure in any critical structural element will be detected before the strength of the element has been reduced below the level necessary to withstand the loading conditions specified in § 25.571(b), thereby allowing timely replacement or repair of the failed elements.
b. Damage-Tolerance Assessment Methodologies. Normally, the damage-tolerance assessment consists of a deterministic evaluation of the design features described in this section.
Sections 6c through 6j below provide guidelines for this approach. In certain specific instances, however, damage-tolerant design might be more realistically assessed by a probabilistic evaluation, employing methods such as risk analysis. Risk analyses are routinely employed in fail-safe evaluations of airplane systems and have occasionally been used where structure and systems are interrelated. These methods can be of particular value for structure consisting of discrete elements, where damage tolerance depends on the ability of the structure to sustain redistributed loads after failures resulting from fatigue, corrosion, or accidental damage. Where considered appropriate on multiple load path structure, a probabilistic analysis may be used if it can be shown that (1) loss of the airplane is extremely improbable, and (2) the statistical data employed in the analysis of similar structure is based on tests or operational experience, or both.
c. Identification of principal structural elements. As defined in this AC, a principal structural element is an element of structure that contributes significantly to the carrying of flight, ground, or pressurization loads and whose integrity is essential in maintaining the overall structural integrity of the airplane. Principal structural elements include all structure susceptible to fatigue cracking, which could contribute to a catastrophic failure. Refer to appendix 5 of this AC for clarification on how this term relates to the terms, fatigue critical structure (FCS) and WFD – susceptible structure. Examples of such elements are as follows:
(1) Wing and empennage.
(a) Control surfaces, slats, flaps, and their mechanical systems and attachments (hinges, tracks, and fittings);
(b) Integrally stiffened plates;
(c) Primary fittings;
(d) Principal splices;
(e) Skin or reinforcement around cutouts or discontinuities;
(f) Skin-stringer combinations;
(g) Spar caps; and
(h) Spar webs.
(2) Fuselage.
(a) Circumferential frames and adjacent skin;
(b) Door frames;
(c) Pilot-window posts;
(d) Pressure bulkheads;
(e) Skin and any single frame or stiffener element around a cutout;
(f) Skin or skin splices, or both, under circumferential loads;
(g) Skin or skin splices, or both, under fore and aft loads;
(h) Skin around a cutout;
(i) Skin and stiffener combinations under fore and aft loads;
(j) Door skins, frames, and latches; and
(k) Window frames.
(3) Landing gear and their attachments.
(4) Engine mounts.
d. Extent of damage. Each particular design should be assessed to establish appropriate damage criteria in relation to inspectability and damage-extension characteristics. In any damage determination, including those involving multiple cracks, it is possible to establish the extent of damage in terms of the following parameters:
detectability with the inspection techniques to be used, 6
the associated, initially detectable crack size, the residual-strength capabilities of the structure, and the likely damage-extension rate.
This determination should consider the expected stress redistribution under the repeated loads expected in service at the expected inspection frequency. Thus, an obvious partial failure could be the extent of the damage for residual-strength assessment, provided that the fatigue cracks will be detectable at a sufficiently early stage of crack development. The following are examples of partial failures that should be considered in the evaluation:
(1) Detectable skin cracks emanating from the edge of structural openings or cutouts;
(2) A detectable circumferential or longitudinal skin crack in the basic fuselage structure;
(3) Complete severance of interior frame elements or stiffeners in addition to a detectable crack in the adjacent skin;
(4) A detectable failure of one element of components in which dual construction is used, such as spar caps, window posts, window or door frames, and skin structure;
(5) A detectable fatigue failure in at least the tension portion of the spar web or similar element; and
(6) The detectable failure of a primary attachment, including a control surface hinge and fitting.
e. Inaccessible areas. Every reasonable effort should be made to ensure inspectability of all structural parts and to qualify those parts under the damage-tolerance provisions (see § 25.611).
f. Testing of principal structural elements. The nature and extent of residualstrength tests on complete structures or on portions of the primary structure depends upon applicable previous design, construction, tests, and service experience with similar structures.
Simulated cracks should be as representative as possible of actual fatigue damage. Where it is not practical to produce actual fatigue cracks, damage can be simulated by cuts made with a fine saw, sharp blade, guillotine, or other suitable means. If saw cuts in primary structure are used to simulate sharp fatigue cracks, sufficient evidence should be available from element tests to indicate equivalent residual strength. In those cases where bolt failure or its equivalent is to be simulated as part of a possible damage configuration in joints or fittings, bolts can be removed to provide that part of the simulation.
g. Identification of locations to be evaluated. The locations of damage to structure for damage-tolerance evaluation should be identified as follows:
(1) Determination of general damage locations. The evaluation would include a determination of the probable locations and modes of damage due to fatigue, corrosion, or accidental damage. Repeated load and static analyses, supported by test evidence and (if available) service experience, would also be incorporated in the evaluation. Special consideration for widespread fatigue damage must be included where the design is such that this type of damage could occur. The location and modes of damage can be determined by analysis or service experience, or by fatigue tests on complete structures or subcomponents. However, tests may be necessary when the basis for analytical prediction is not reliable, such as for complex components. If less than the complete structure is tested, ensure that the internal loads and boundary conditions are valid.
(a) If a determination is made by analysis, take into account factors such as the following:
1 Strain data on undamaged structure to establish points of high-stress concentration as well as the magnitude of the concentration;
2 Locations where permanent deformation occurred in static tests;
3 Locations of potential fatigue damage identified by fatigue analysis;
and 4 Design details that service experience, of similarly designed components, indicates are prone to fatigue or other damage.
(b) In addition, the areas of probable damage from other sources, such as severe corrosion or accidental damage, should be determined from a review of the design and past service experience.
(2) Selection of critical damage areas. The process of actually locating where damage should be simulated in principal structural elements (identified in section 6d of this AC) should take into account factors such as the following:
(a) Review analysis to locate areas of maximum stress and low margin of safety;
(b) Select locations in an element where the stresses in adjacent elements would be the maximum with the damage present;
(c) Select partial-fracture locatio ns in an element where high stress concentrations are present in the residual structure; and
(d) Select locations where detection would be difficult.
h. Damage-tolerance analysis and tests.
(1) Analysis, supported by test evidence, should determine that:
(a) The structure with the extent of damage established for residualstrength evaluation can withstand the specified design-limit loads (considered as ultimate loads);
and
(b) The damage-growth rate under the repeated loads expected in service – between the time the damage becomes initially detectable and the time the extent of damage reaches the value for residual-strength evaluation – provides a practical basis for development of the inspection program and procedures described in section 6j of this AC.
(2) The repeated loads should be as defined in the loading, temperature, and humidity spectra. The loading conditions should take into account the effects of structural flexibility and rate of loading where they are significant.
(3) The damage-tolerance characteristics can be shown analytically by reliable or conservative methods, such as the following:
(a) Demonstrating quantitative relationships with structure already verified as damage tolerant;
(b) Demonstrating that the damage would be detected before it reaches the value for residual-strength evaluation; or
(c) Demonstrating that the repeated loads and limit-load stresses do not exceed those of previously verified designs of similar configuration, materials, and inspectability.
(4) The maximum extent of immediately obvious damage from discrete sources should be determined, and the remaining structure shown to have static strength for the maximum load (considered as ultimate load) expected during the remainder of the flight.
Normally, this would be an analytical assessment. In the case of uncontained engine failure, the fragments and paths to be considered should be consistent with those used in showing compliance with § 25.903(d)(1) and with typical damage experienced in service. (See AC 20- 128, “Design Considerations for Minimizing Hazards Caused by Uncontained Turbine Engine and Auxiliary Power Unit Rotor and Fan Blade Failures.”)
i. Inspection.
(1) Detection of damage before it becomes critical is the ultimate control in ensuring the damage-tolerance characteristics of the structure. For this reason, Amendment 25-54 revised § 25.571 to require that the applicant (1) establish inspections or other procedures, as necessary, to prevent catastrophic failure from accidental, environmental, or fatigue damage, and (2) include those inspections and procedures in the ALS of the ICA, as required by § 25.1529 (see also appendix H to part 25).
(2) Due to the complex interactions of the many parameters affecting damage tolerance – such as operating practices, environmental effects, load sequence on crack growth, and variations in inspection methods – take into account related operational experience in 9
establishing inspection procedures. Additionally, give careful consideration to the practical nature of inspection procedures.
(3) Comparison with past successful practice is the primary means of substantiating inspections or other procedures for accidental and environmental damage. For a new-model transport category airplane, the Maintenance Review Board generally conducts such comparison to substantiate inspections or other actions using the Air Transport Association of America, Inc. (ATA) Maintenance Steering Group’s MSG-3 or other accepted version of the “Operator/Manufacturer Scheduled Maintenance Development” procedures. If this process is used, the required maintenance actions for accidental and environmental damage must be documented in the Maintenance Review Board Report for the airplane model and must be complete not later than when the first airplane enters service. These inspections or other procedures, as necessary to prevent catastrophic failure of the airplane, must be included in the Airworthiness Limitations section of the ICA. Alternatively, the applicant may reference, in the ALS of the ICA, the maintenance documents that contain those tasks. The ALS should also contain reference to any corrosion prevention and control program (CPCP) developed to maintain corrosion to “Level 1” or better for that airplane model. “Level 1” corrosion is damage occurring between successive inspections that is local and can be reworked/blended-out within allowable limits as defined by the manufacturer’s service information, such as structural-repair manuals and service bulletins.
(4) In some cases, the experience of an operator may indicate that different inspections or other procedures are appropriate for that operator. Title 14 CFR 43.16 and 14 CFR 91.403(c) provide a means for FAA approval of alternatives to the airworthiness limitations. The FAA will evaluate such proposed alternatives, using the methods described in this AC or other acceptable methods proposed by the type-certificate holder and operators, to ensure that the objectives of 14 CFR 25.571 continue to be met.
j. Threshold for Inspections.
(1) Where it can be shown by observation, analysis, and/or test that a load path failure in multiple load path “fail-safe” structure or partial failure in crack-arrest “fail-safe” structure will be detected and repaired during normal maintenance, inspection, or operation of an airplane prior to failure of the remaining structure, the thresholds can be established using either:
(a) Fatigue analysis and tests with an appropriate scatter factor; or
(b) Slow-crack-growth analyses and tests, based on appropriate initial manufacturing damage.
(2) For single load path structure and for multiple load path and crack-arrest “fail-safe” structure – where it cannot be demonstrated that load-path failure, partial failure, or crack arrest will be detected and repaired during the normal maintenance, inspection, or operation of an airplane prior to failure of the remaining structure – the thresholds should be established based on crack-growth analyses and/or tests, assuming the structure contains an initial flaw of the maximum probable size that could exist as a result of manufacturing- or service-induced damage.
7. Establishing an LOV.
a. Structural-maintenance program.
(1) If an airplane is properly maintained, theoretically it could be operated indefinitely. However, it should be noted that structural-maint enance tasks for an airplane are not constant with time. Tasks typically are added to the maintenance program as the airplane ages. It is reasonable to expect, then, that conf idence in the effectiveness of current structuralmaintenance tasks may not, at some future point, be sufficient for continued operation.
Maintenance tasks for a particular airplane can only be determined based on what is known about that airplane model at any given time; from analyses, tests, service experience, and teardown inspections. Widespread fatigue damage is of particular concern because inspection methods cannot be relied on solely to ensure the continued airworthiness of airplanes indefinitely. To preven t WFD from occurring, the structure occasionally must be modified or replaced. Establishing a ll the replacements and modifications required to operate the airplane indefinitely is an unbounded problem. This problem is solved by establishing limit of validity of the engineering data that supports the structural-maintenance program. All necessary modifications and replacements are required to be established to ensure continued airworthiness relative to WFD up to the LOV. To operate beyond the LOV, the full-scale fatigue-test evidence and the structural maintenance program must be re-evaluated to determine if additional modifications or replacements are required. See paragraph 7.g for the steps to extend the LOV.
b. Widespread Fatigue Damage.
(1) Structural fatigue damage is progressive. It begins as minute cracks, and those cracks grow under repeated stresses. It can initiate as a result of normal operational conditions and design attributes, or from isolated incidents, such as material defects, poor fabrication quality, or corrosion pits, dings, or scratches. Fatigue damage can occur locally, in small areas or structural design details, or globally. Global fatigue damage is a general degradation of large areas of structure that share similar structural details and stress levels.
Global damage may occur in a large structural element, such as a single rivet line of a lap splice joining two large skin panels (multiple site damage). Or, it may be found in multiple elements, such as adjacent frames or stringers (multiple element damage). Multiple site damage and multiple element damage cracks are typically too small, initially, to be reliably detected with normal inspection methods. Without intervention, these cracks will grow, and eventually compromise the structural integrity of the airplane, in a condition known as widespread fatigue damage. Widespread fatigue damage is increasingly likely as an airplane ages, and is certain if an airplane is operated long enough without any intervention.
c. Steps for Establishing an LOV.
(1) Under § 25.571, persons applying for type certificates (TC) must establish an LOV for each new type design. For new airplane models, or design changes to existing models that are added to the TC, the type-certificate holder must either establish an LOV for the new or design-changed model, or show that the LOV for the originally certificated model can also be applied to the new or design-changed model. The LOV is the period of time (in flight cycles, flight hours, or both) up to which it has been demonstrated that widespread fatigue damage is 11
unlikely to occur in an airplane’s structure by virtue of its inherent design characteristics and any required maintenance actions. An airplane may not operate beyond the LOV. To support establishment of the LOV, the applicant must demonstrate, by test evidence and analysis at a minimum and, if available, service experience, or service experience and teardown-inspection results of high-time airplanes, that WFD is unlikely to occur in that airplane up to the LOV. An LOV applies to an airplane structural configuration common to a fleet.
(2) The process for establishing an LOV involves four steps— Identifying a “candidate LOV.”
Identifying WFD-susceptible structure.
Performing a WFD evaluation of all susceptible structure.
Finalizing the LOV and establishing necessary maintenance actions.
(a) Step 1 – Candidate LOV. Any LOV can be valid as long as it has been demonstrated that the airplane model will be free from WFD up to the LOV based on the airplane’s inherent fatigue characteristics and any required maintenance actions. Early in the certification process, applicants typically establish design service goals or their equivalent and set a design service objective to have structure remain relatively free from cracking, up to the design service goal. A recommended approach sets the “candidate LOV” equal to the design service goal. The final LOV would depend on bot h how well that design objective was met and the applicant’s consideration of the economic impact of maintenance actions required to preclude WFD up to the final LOV.
(b) Step 2 – Identify WFD-Susceptible Structure. The applicant should identify the structure that is susceptible to WFD to support post-fatigue-test teardown inspections or residual-strength testing necessary to demonstrate that WFD will not occur in the airplane structure up to the LOV. Appendix 5 of AC 120-104 provides examples and illustrations of structure where multiple site damage or multiple element damage has been documented. The list in appendix 5 is not meant to be inclusive of all structure that might be susceptible on any given airplane model and it should only be used for general guidance. It should not be used to exclude any particular structure. The applicant should do the following when developing the list of structure susceptible to WFD:
1 Establish criteria that could be used for identifying what structure is susceptible to WFD based on the definitions (see appendix 1) of multiple site damage, multiple element damage, and WFD. For example, structural details and elements that are repeated over large areas and operate at the same stress levels are obvious candidates. The criteria should be part of the applicant’s compliance data.
2 Provide supporting rationale for including and excluding specific structural areas. This should be part of the applicant’s compliance data.