7.2Established Approaches for Specific Fracture Critical Hardware Types

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results in loss of structural margin or the need for operational modifications, or loss of critical system function) may be proposed for acceptance without damage tolerance assessment (in combination with other activities) by developing an alternative approach as required in FCR [26], section 10 of this NASA Technical Standard.

Equivalence means that damage tolerance life analysis or test requirements in sections 7.2.1 or 7.2.2 in this NASA Technical Standard are also applied in modified form for a vessel meeting section 7.2.3 in this NASA Technical Standard. Equivalence does not mean other types of assessment, such as fatigue calculations or cycle test, can be substituted for the damage tolerance methodology detailed in sections 7.2.1 and 7.2.2 in this NASA Technical Standard.

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• (1) All occurrences of "maximum expected operating pressure" and "MEOP" are substituted with "maximum design pressure" and "MDP" as terms in this NASA Technical Standard in section 3.2.

• (2) The word "nominal" is replaced with the word "average" in all ANSI/AIAA S-080-1998 sections except 4.7.2.

• (3) All occurrences of the term “service life” have the meaning defined in this NASA Technical Standard in section 3.2 for “service life.”

• c. The ANSI/AIAA S-080-1998 requirements in section 5.1, Approach A, Path 2, as detailed in section 5.1.2 of that document are followed for all the metallic pressure vessels addressed by that section with the modifications specified in this section of this NASA Technical Standard.

• d. ANSI/AIAA S-080-1998 section 4.2.7 safe-life requirements are met with the following modifications:

• (1) The safe-life assessment analysis and test assessments are to encompass and represent the worst-case flaw location, shape, aspect ratio, and orientation.

• (2) The process for selecting the worst-case flaw location, shape, aspect ratio, and orientation is based on vessel stress/strain response, material strength, and crack growth properties and documented in the analysis report.

• (3) The assessment determining the worst-case flaw location, shape, aspect ratio, and orientation includes all regions of the pressure vessel, including the boss and any internal and external attachments.

• (4) The safe-life assessment analysis and test loading spectra are to include all loadings experienced during the service life, including those specified in this NASA Technical Standard in section 7.3.1, unless the RFCB approves the exclusion of specific loadings as insignificant for a component assessment. For example, with approval of the RFCB, service life loadings that affect the safe-life of a particular region of the vessel by less than 5 percent may be excluded from the safe-life assessment of these regions.

• (5) The assessments are to show that all safe-life requirements are met for the entire mission service life.

The mission service life includes all of the hardware activities included in the hardware mission as defined in NPR 7120.5, for the duration of the service life as defined in section 3.2 in this NASA Technical Standard. If the mission service life includes periodic "depot" intervals (opportunities for inspection) with fully

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qualified screening inspections that ensure that acceptable hardware has sufficient life, including the service life factor, to reach the next "depot" evaluation, this "depot" interval-based service-life approach may be proposed as an alternative approach by meeting the requirements in section 10 of this NASA Technical Standard.

• e. If the AIAA S-080-1998 section 4.2.7 analysis option to show safe-life is planned, apply the following modifications to the requirements:
• (1) Obtain pre-approval by the RFCB for all crack growth computer analysis programs other than NASGRO®.
• (2) If the analysis ability to simulate crack growth is invalidated by plasticity or other effects, the assessment is performed by test.
• (3) If NASGRO® is used, either set Bk to zero, or set Bk such that the stress intensity factor for the part thickness is less than or equal to the critical stress intensity value with approval of the Technical Authority or the RFCB.
• (4) Establish that the assessed parts survive 4 lifetimes without failure (hazardous leak or fracture instability) by analyses that assess all applicable effects causing crack growth as a result of cyclic loadings.
• A. If the loading sequence of high/low loads is unknown, then damage tolerance analysis is to show that the stress intensity factor at limit load is less than the critical stress intensity factor or residual strength at the end of 4 lifetimes.
• B. If the service lifetime is a single event or the fatigue crack growth is small relative to the critical crack size (initial and critical cracks are of similar size), the analysis is to establish one of the following:
• i. Reserve capability against fracture by meeting either a lower bound critical stress intensity factor or residual strength at the end of 4 lifetimes.
• ii. A factor of 1.4 on critical stress intensity factor or residual strength after 1 lifetime.

Assessments of metallic alloys that are susceptible to crack growth related to SLC or EAC during the service life are addressed in item (6) below.

• (5) Use critical stress intensity factor and cyclic threshold stress intensity range (ΔKth) values that are less than or equal to the average values.
• (6) For metallic alloys susceptible to EAC or SLC or both, satisfy all of the following:

Assessments of metallic alloys that are susceptible to crack growth related to SLC or EAC during the service life are addressed in item 6 below.

• A. If the loading sequence of high/low loads is unknown, then damage tolerance analysis is to show that the stress intensity factor at limit load is less than the critical stress intensity factor or that the part applied load does not exceed the residual strength at the end of 4 lifetimes.

• B. If the service lifetime is a single event or the amount of fatigue crack growth is small relative to the critical crack size for unstable crack growth, the analysis is to show reserve capability against fracture by meeting either of the following:

• i. A lower bound critical stress intensity factor or residual strength at the end of 4 lifetimes.

• ii. A factor of 1.4 on critical stress intensity factor or residual strength after 1 lifetime.

• (5) Use critical stress intensity factor and cyclic threshold stress intensity range (ΔKth) values that are less than or equal to the average values.

• (6) For metallic alloys susceptible to EAC or SLC or both, satisfy all of the following:

• A. Use the lower bound value of stress intensity factor threshold for assessment of EAC (KEAC or KIEAC as appropriate), and SLC if the material exhibits these behaviors in the application conditions.

• B. Show that the applied stress intensity factor related to the largest service load is smaller than the lower bound stress intensity factor thresholds determined in item A above at the end of 4 lifetimes.

• f. When performing proof testing in accordance with ANSI/AIAA S-081-2000 section 5.1.2, the duration of the proof test loading is minimized while also meeting the requirement to verify the pressure stability.

• g. When performing assessment to show safe-life by test for non-linear response of the metal liner in accordance with ANSI/AIAA S-081-2000 section 5.2.1, apply the following items:

• (1) The testing approach and rationale are subject to both of the following:

• A. Provided to the RFCB for approval before implementation.

• B. Documented in the FCP.

• A. Use the lower bound value of stress intensity factor threshold for assessment of EAC (KEAC or KIEAC as appropriate) and SLC if the material exhibits these behaviors in the application conditions.

• B. Show that the applied stress intensity factor related to the largest service load is smaller than the lower bound stress intensity factor thresholds determined in item A above at the end of 4 lifetimes.

• f. When performing proof testing in accordance with ANSI/AIAA S-080-1998 sections 4.2.7, 4.6.4, and/or 5.1.2.4, the duration of the proof test loading is minimized while also meeting the requirement to verify the pressure stability.

• g. If the AIAA S-080-1998 section 4.2.7 testing option to show safe-life is planned, requirements are to include the following items:

• (1) The testing approach and rationale are subject to both of the following:

• A. RFCB approval before implementation.

• B. Documentation in the FCP.

• (2) The testing is to show that the hardware meets the damage tolerance lifetime and failure condition requirements in ANSI/AIAA S-080-1998 as modified in this NASA Technical Standard for initial flaws in the worst location, aspect ratio, and orientation in conditions that account for the service environments.

• (3) Testing reports showing that the testing objectives have been achieved are documented in accordance with section 9.1 in this NASA Technical Standard and cited in the FCSR.

• h. The ANSI/AIAA S-080-1998 section 5.1.2.6 Special Provision is not allowed.

Pressure vessels as defined by NASA are always fracture critical.

• i. Vessels with crack-like flaws that are induced during the manufacturing process are not accepted as flight hardware unless a process for remediation repair has been established and the Technical Authority approves the part and process.

Refer to section 8.1.5 of this NASA Technical Standard for further requirements and guidance.

• j. The ANSI/AIAA S-080-1998 requirements are subject to the following:
• (1) Quality assurance in section 4.6 of that document is supplemented by requirements in section 8 (and its subsections) in this NASA Technical Standard.

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• (2) If there is a conflict with ANSI/AIAA S-080-1998, the ANSI/AIAA S-080-1998 requirements for quality assurance in section 4.6 of that document are superseded by requirements in section 8 (and its subsections) of this NASA Technical Standard.
• k. The ANSI/AIAA S-080-1998 requirements for fracture critical part documentation and reporting are subject to the following:
• (1) Supplemented by requirements in section 9 (and its subsections) of this NASA Technical Standard.
• (2) If there is a conflict with ANSI/AIAA S-080-1998, the ANSI/AIAA S-080-1998 requirements for fracture critical part documentation and reporting of that document are superseded by requirements in section 9 (and its subsections) of this NASA Technical Standard.
• (3) The ANSI/AIAA S-080-1998 section 4.2.5 required stress analysis report and the section 4.2.7 safe-life analysis report are provided as part of the FCSR documentation.

Note that ANSI/AIAA S-080-1998 also addresses other hardware types, but only the metallic pressure vessel requirements as tailored in this section are applicable for this NASA Technical Standard.

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• a. The ANSI/AIAA S-081-2000 requirements are followed for the assessment and qualification of all the composite overwrapped vessels and composite overwrapped pressurized fluid containers with metallic liners addressed by this section, regardless of the vessel fluid pressure, energy, or hazardous nature, with the modifications specified in this section in this NASA Technical Standard.
• b. The damage tolerance assessment approach is described in the FCP in accordance with section 4.1 [FCR 1] in this NASA Technical Standard.
• c. All occurrences of the following terms in ANSI/AIAA S-081-2000 are replaced with the terms having meanings as specified below:
• (1) All occurrences of "maximum expected operating pressure" and "MEOP" are substituted with "maximum design pressure" and "MDP" as defined in this NASA Technical Standard in section 3.2.
• (2) The word "nominal" is replaced with the word "average" in all ANSI/AIAA S-081-2000 sections.
• (3) All occurrences of the term "service life" are to have the meaning defined in this NASA Technical Standard in section 3.2 for "service life."
• d. ANSI/AIAA S-081-2000 section 4.2 requirements are met with the following modifications:
• (1) In ANSI/AIAA S-081-2000 section 4.2, the damage tolerance, i.e., safe-life, approach (b) is the only acceptable approach.
• (2) The ANSI/AIAA S-081-2000 section 4.2.7 safe-life requirements are met with the following modifications:
• A. The safe-life assessment analysis and test assessments are to encompass and represent the worst-case flaw location, shape, aspect ratio, and orientation.
• B. The process for selecting the worst-case location, shape, aspect ratio, and orientation is based on liner stress/strain response and material strength and crack growth properties and documented in the analysis report.
• C. The assessment determining the worst-case location, shape, aspect ratio, and orientation includes all regions of the liner and boss, including the shear region of the boss and any internal and external attachments.
• D. The safe-life assessment analysis and test loading spectra are to include all loadings experienced during the service life, including those specified in this NASA Technical Standard in section 7.3.1, unless the RFCB approves the

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exclusion of specific loadings as insignificant for a component assessment.

For example, with approval of the RFCB, service life loadings that affect the safe-life of a particular region of the liner, boss, or shear region of the boss by less than 5 percent may be excluded from the safe-life assessment of these regions.

• E. The assessments are to show all safe-life requirements are met for the entire mission service life.

The mission service life includes all of the hardware activities included in the hardware mission as defined in NPR 7120.5, for the duration of the service life as defined in section 3.2 in this NASA Technical Standard. If the mission service life includes periodic "depot" intervals (opportunity for inspection) with fully qualified screening inspections that ensure acceptable hardware has sufficient life, including the service life factor, to reach the next "depot" evaluation, this "depot" interval-based service-life approach may be proposed as an alternative approach by meeting the requirements in section 10 in this NASA Technical Standard.

• (3) Autofrettage is included in the service life unless liner NDE is performed after autofrettage.
• (4) The assessment of crack growth related to the autofrettage cycle is determined by test in accordance with ANSI/AIAA S-081-2000 section 5.2.1, unless prior approval is provided by the RFCB for an analytical approach.
• e. When performing analysis to show safe life for linearly responding portions of the metal liner in accordance with ANSI/AIAA S-081-2000 section 4.2.7, apply the following modifications:
• (1) Obtain pre-approval from the RFCB for all crack growth computer analysis programs other than NASGRO®.
• (2) If the analysis ability to simulate crack growth is invalidated by plasticity or other effects, the assessment is performed by test.
• (3) If NASGRO® is used, either set Bk to zero, or set Bk such that the stress intensity factor for the part thickness is less than or equal to the critical stress intensity value with approval of the Technical Authority or the RFCB.
• (4) The analysis shows that the parts survive 4 service lives without failure by assessments that address all applicable effects causing crack growth as a result of cyclic loading, using the following criteria:

Controlled by: NASA Assessments of metallic alloys that are susceptible to crack growth related to SLC or EAC during the service life are addressed in item 6 below.

A. If the loading sequence of high/low loads is unknown, then damage tolerance analysis is to show that the stress intensity factor at limit load is less than the critical stress intensity factor or that the part applied load does not exceed the residual strength at the end of 4 lifetimes.

B. If the service lifetime is a single event or the amount of fatigue crack growth is small relative to the critical crack size for unstable crack growth, the analysis is to show reserve capability against fracture by meeting either of the following:

i. A lower bound critical stress intensity factor or residual strength at the end of 4 lifetimes.

ii. A factor of 1.4 on critical stress intensity factor or residual strength after 1 lifetime.

(5) Use critical stress intensity factor and cyclic threshold stress intensity range (ΔKth) values that are less than or equal to the average values.

(6) For metallic alloys susceptible to EAC or SLC or both, satisfy all of the following:

A. Use the lower bound value of stress intensity factor threshold for assessment of EAC (KEAC or KIEAC as appropriate), and SLC if the material exhibits these behaviors in the application conditions.

B. Show that the applied stress intensity factor related to the largest service load is smaller than the lower bound stress intensity factor thresholds determined in item A above at the end of 4 lifetimes.

f. When performing proof testing in accordance with ANSI/AIAA S-081-2000 section 5.1.2, the duration of the proof test loading is minimized while also meeting the requirement to verify the pressure stability.

g. When performing assessment to show safe-life by test for non-linear response of the metal liner in accordance with ANSI/AIAA S-081-2000 section 5.2.1, apply the following items:

(1) The testing approach and rationale are subject to both of the following:

A. Provided to the RFCB for approval before implementation. B. Documented in the FCP.

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• (2) The testing is to show that the hardware meets the damage tolerance lifetime and failure condition requirements in ANSI/AIAA S-081-2000 as modified in this NASA Technical Standard for initial flaws in the worst location and orientation in conditions that account for the service environments.
• (3) Testing reports showing that the testing objectives have been achieved and are documented in accordance with section 9.1 of this NASA Technical Standard and cited in the FCSR.
• h. ANSI/AIAA S-081-2000 section 4.2.10 damage control requirements are to include the section 4.2.10.2.1 protective cover approach with the following additional requirements:
• (1) The covers are required regardless of the COPV burst factor, wall thickness, hazardous or nonhazardous nature of the fluid, or energy content.
• (2) If the vessel is exposed to risk of damage during any parts of the service life where the initially applied covers are not present, additional damage controls are selected from the options in ANSI/AIAA S-081-2000 section 4.2.10.

The additional damage controls referenced in item h.2 above may be needed if protective covers are removed before launch and the vessel has risk of damage during the remainder of its service life or if there are different risks to the vessel during its service life because of environments or service loadings that also need to be addressed by damage controls. All options, separately or in combination, may be used, including specialized covers for flight conditions.

• i. Apply the following items if the composite overwrap is constrained by external structure or if it is part of a load path supporting the COPV for service life loads other than pressure loads:
• (1) Perform an assessment validated by testing that shows the overwrap with the external structure loads meets all strength, fatigue, and life requirements in ANSI/AIAA S-081-2000.
• (2) The assessment is to include effects of damage conditions that are not screened by the protections imposed in accordance with ANSI/AIAA S-081-2000 section 4.2.10 with the modifications in this section of this NASA Technical Standard.
• j. Vessels with crack-like flaws in the metal liner that are induced during the manufacturing process are not accepted as flight hardware unless a process for remediation repair has been established and the Technical Authority approves the part and process.

Refer to section 8.1.5 in this NASA Technical Standard for further requirements and guidance.

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• k. Damage in other regions of the vessel may be repaired with an established, proven process if approved by the Technical Authority.

Refer to section 8.1.5 of this NASA Technical Standard for further requirements and guidance. This pertains to the repair of small manufacturing or cosmetic defects in the composite. There are no acceptable established processes for repairing impact damage to the composite overwrap. Accidental impacts that do not leave obvious visible damage indications are to be logged, the impact site assessed by qualified inspectors, and the hardware approved for use by the Technical Authority.

1. The ANSI/AIAA S-081-2000 requirements for quality assurance in section 4.5 in that document are supplemented and superseded by requirements in section 8 (and its subsections) in this NASA Technical Standard.

• m. The ANSI/AIAA S-081-2000 requirements for fracture critical part documentation and reporting are subject to the following:
• (1) Requirements in section 9 (and its subsections) of this NASA Technical Standard.
• (2) If there is a conflict with ANSI/AIAA S-081-2000, the requirements are superseded by requirements in section 9 (and its subsections) of this NASA Technical Standard.
• (3) The ANSI/AIAA S-081-2000 section 4.2.7 safe-life and analysis reports and the 4.2.10 Mechanical Damage Control Plan (MDCP) are provided as part of the FCSR documentation.

The entity responsible for delivery of the MDCP (NASA, prime contractor, or other subcontractors) determines who develops the MDCP, which is subject to RFQB approval.

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• B. A COPV or composite overwrapped pressurized fluid container that meets the requirements of section 7.2.2 in this NASA Technical Standard.
• (2) Describe the proposed approach that satisfies applicable requirements in items b, c, or d below in this section and the requirements in sections 8 and 9 in this NASA Technical Standard.
• (3) Receive RFCB approval before implementing the proposed approach.

A rationale is required because detailed requirements for the approach have to be developed and documented in the FCP that satisfy the applicable requirements in b or c below and the guidance in this section, and the RFCB has to review and approval the proposed detailed approach. This presents a significant effort for the developer of the FCP and for the RFCB reviews.

• b. The development approach is satisfied by comparison to requirements in sections 7.2.1 in this NASA Technical Standard for metallic pressure vessels or 7.2.2 in this NASA Technical Standard for COPVs and composite overwrapped pressurized fluid containers. The approach is to be equivalent to or an extension of all the requirements, including establishing that damage tolerance life is achieved without failure or leakage of the fluid, and provides equivalent risk mitigation of a catastrophic failure caused by flaws.
• c. The proposed FCP approach for damage tolerance assessment of a fracture critical pressure vessel or pressurized fluid container that is all composite or has a non-metallic, i.e., an elastomeric, liner or other non-metallic components is to meet the general approach for fracture critical composite hardware in section 7.4 in this NASA Technical Standard and show that the damage tolerance required life is achieved without failure or leakage of the fluid.
• d. The proposed FCP approach for damage tolerance assessment of a fracture critical all-metal pressurized fluid container is to meet the general approach for fracture critical metallic hardware in section 7.3 in this NASA Technical Standard and show that the damage tolerance life is achieved without failure or leakage of the fluid.

Note that if a fracture critical metallic "pressurized fluid container" is planned with attributes close to the definition of a pressure vessel, it may be advantageous to push it into the pressure vessel category to minimize later impacts as the project matures in case the initial design attributes increase. There are currently no predefined approaches for pressure vessels or pressurized fluid containers that are qualified under a different code/standard than ANSI/AIAA S-080 or ANSI/AIAA S-081, such as the ASME Boiler and Pressure Vessel Code, Section VIII, Divisions 1 or 2, or the United States Department of Transportation Code of Federal Regulations Title 49, Transportation. These codes/standards do not impose the structural integrity activities needed for damage tolerance that are specified in ANSI/AIAA S-080 and ANSI/AIAA S-081. As a result, the approaches used by these codes/standards to certify vessels do not facilitate meeting damage tolerance requirements as required in this NASA Technical Standard. Equivalence means that

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damage tolerance life analysis or test requirements in sections 7.2.1 or 7.2.2 in the NASA Technical Standard are also applied in modified form for a vessel meeting section 7.2.3 in this NASA Technical Standard. Equivalence does not mean other types of assessment, such as fatigue calculations or cycle test, can be substituted for the damage tolerance methodology detailed in sections 7.2.1 and 7.2.2 in this NASA Technical Standard. Use of these codes/standards in combination with other activities may be proposed, however, as an alternative approach as described in [FCR 26] in section 10 in this NASA Technical Standard.

In addition, other pressure vessels and pressurized fluid containers may be developed that are not addressed by existing codes or standards. Examples may include composite pressure vessels/containers without a metal liner or rubber-lined composite pressure vessels/containers. For these other fracture critical vessels/containers, a unique approach is developed and proposed in the FCP that establishes equivalent methods of addressing material, structural, qualification, acceptance, and related aspects such as those in the ANSI/AIAA S-080 or ANSI/AIAA S-081 standards to support the damage tolerance assessment. Equivalence with the AIAA pressure vessel standards may include assessments and testing that include materials aspects, loadings, stress analysis, strength, environment effects, stiffness, thermal response, life, quality assurance, repairs, NDE requirements, acceptance processes including proof and leakage testing, damage tolerance control plans, and damage tolerance assessments by analysis and/or testing, and documentation. However, use of analytical techniques to establish damage tolerance is generally considered insufficiently developed for composite pressure vessels. For all-composite pressure vessels, the approaches described for fracture critical composite hardware in section 7.4 in the NASA Technical Standard should be incorporated, in addition to applicable equivalent requirements in ANSI/AIAA S-081.

For other types of vessels/containers, it should also be noted that, in addition to the section 9 documentation in this NASA Technical Standard showing the approach proposed in the FCP has been met, section 9.1.3.1 in this NASA Technical Standard requires providing supporting detailed technical information to the RFCB upon request, including drawings, material and processing data, detailed stress analysis, and damage tolerance analyses that are needed to support the damage tolerance assessment.

Early involvement with the RFCB is suggested for any vessels/containers to be assessed by this section.

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• a. Apply the following items (1) through (6) to parts where the only load of significance is related to pressure:
• (1) The metallic material is not susceptible to crack extension related to EAC or SLC.
• (2) Perform 100 percent inspection of all fusion joints in fracture critical pressure components using a qualified NDE method after proof test to inspect for the presence of unacceptable lack of penetration or other unacceptable conditions both on the surface and within the fusion joint.
• (3) Reject any type of flaw indication in the final product that does not meet specification requirements.

NDE rejection indicates the need for formal review and part disposition.

• (4) Proof test lines, fittings, joints, and other pressurized components or parts to a minimum of 1.5 times the MDP during individual acceptance or at the system level.
• (5) An ECF less than 1.0 is not allowed without prior approval by the RFCB.
• (6) Obtain RFCB approval that the part is manufactured using processes that have been established by reliability or by inspections of many similar parts to be extremely unlikely to produce parts with a flaw exceeding process specifications.

For loading (stresses) to be considered pressure dominant, all other loads (stresses) should be no greater than 20 percent of the pressure loads (stresses).

• b. Satisfy section 7.3 in this NASA Technical Standard for parts that do not meet the criterion in 7.2.4.a. Item 7.2.4.a is intended for hardware designed to carry primarily pressure loads. This hardware is designed with appropriate supports, brackets, or relief loops such that they are not subject to significant structural loads. Typically, these parts are produced under process control in large quantities, are identical parts, and are subjected to NDE and qualification testing to ensure the parts are reliable and present a low risk of containing detectable flaws that result in crack growth.

Note that additively manufactured parts are not eligible for this category according to NASA-STD-6030 AMR-15.

Pressurized components may have high pressures and energies, but this type of hardware is subject to high factors of safety imposed by other standards such as NASA-STD-5001. NASA-STD-5001 also requires implementation of AIAA S-080, which has a leak test requirement.

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Standard NDE is acceptable.

• (1) Welded regions where proof testing adequately screens for flaws are subject to the following:
• A. Perform post-proof NDE (surface and volumetric) of all welded regions for the first flight article (as a minimum).
• B. Also perform post-proof NDE of all affected weld regions (including those that are adequately screened for flaws by proof test) subjected to significant process, material, or vendor changes for the first flight article incorporating the significant changes.
• (2) For welded regions where proof testing does not adequately screen for flaws, perform post-proof NDE (surface and volumetric) of all welded regions for all flight articles.
• (3) All weld intersections, weld repair regions, and weld transition regions, including friction plug pull weld regions, are to receive post-proof NDE (surface and volumetric) for all flight articles.
• d. For composite or bonded pressurized structures, provide the damage tolerance approach and rationale to the RFCB for approval before implementation.

For composite or bonded pressurized structures, the requirements in section 7.4 in this NASA Technical Standard are a good starting point as a fracture control approach but will need enhancement to provide adequate protection against catastrophic hazard.

• e. For composite or bonded pressurized structures, perform post-proof NDE as described in section 8.1.2 in this NASA Technical Standard.

Proof tests are usually performed in the operational environment, or the test levels are adjusted via an ECF.

Proof test levels (factors) are defined either by structural requirements or by those developed to provide flaw screening (section 8.1.3 [FCR 18] in this NASA Technical Standard). Section 8.2 in this NASA Technical Standard requires load history traceability for all fracture critical parts. The use of pressurized structures should be monitored with documentation of the operational history to ensure that certification is not invalidated.

The proof test factor for these structures is a minimum of 1.05 in accordance with NASA-STD-5001. This may result in a high stress during proof and possible growth of large flaws in the structure during the proof test. In accordance with the guidance in section 8.1.3 in this NASA Technical Standard, the flaw size used in the life assessment of these structures in

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regions where the proof test is used for flaw screening needs to adequately account for possible flaw growth during the proof test (typically established by laboratory damage tolerance tests).

Although it may be difficult to obtain adequate flaw screening for all welded regions via a proof pressure test because of external vehicle loads, the proof test is designed to provide as much flaw screening for welds as is practical.

Flaw screening for the entire fracture critical structure is required in accordance with section 8 in this NASA Technical Standard. Pre-proof NDE is highly recommended to protect high-value structures and facilities.

Damage tolerant assessment considers the worst-case allowed weld joint peaking and mismatch effects (metallic structures) and residual stress effects (either by analysis or included as a part of material test data) for pressurized structures.

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• (1) Fasteners are fabricated from a metal with high resistance to stress corrosion cracking, as defined in MSFC-STD-3029.
• (2) Fasteners are fabricated, procured, and inspected in accordance with NASA-STD-8739.14, and an equivalent military standard, NAS, proprietary, or commercial aerospace specification approved by the RFCB.
• (3) The fastened joint complies with NASA-STD-5020 without joint separation in the nominal configuration.
• (4) Fasteners have rolled threads and are assessed to demonstrate they meet the fatigue requirements in NASA-STD-5001.
• (5) Fasteners manufactured from titanium alloys require additional coordination with the RFCB for approval.

Titanium alloys, such as Ti-6Al-4V (including annealed and STA conditions), cp-Ti, and other titanium alloys, have potential generic EAC or SLC failure modes that are to be addressed in the assessment with test data from flawed fasteners in the applicable service life environments.

• (6) The fasteners are not made from a low fracture toughness alloy, as defined in section 3.2 in this NASA Technical Standard.

• (7) Fasteners are not reworked or custom made unless the application is approved by the RFCB.

• b. Include preload and its effect on flaws and cyclic stresses in the damage tolerance assessment.

• c. Inspect all fracture critical fasteners by the eddy current NDE technique or use proof testing to screen for flaws.

• d. Assume a flaw in the most critical location of a size consistent with NDE sensitivity or proof-test level in the damage tolerance analysis. General NDE flaw sizes are given in NASA-STD-5009, Nondestructive Evaluation Requirements for Fracture Critical Metallic Components; but for specific guidelines on eddy current methodology, PRC-6509, Process Specification for Eddy Current Inspection, can be used as a reference.

• e. Proof-load test inserts used in conjunction with fracture critical fasteners to a minimum factor of 1.2 after installation.

Controlled by: NASA

(1) Fasteners are fabricated from a metal with high resistance to stress corrosion cracking, as defined in MSFC-STD-3029.

(2) Fasteners are fabricated, procured, and inspected in accordance with NASA-STD- 8739.14, and an equivalent military standard, NAS, proprietary, or commercial aerospace specification approved by the RFCB.

(3) The fastened joint complies with NASA-STD-5020 without joint separation in the nominal configuration.

(4) Fasteners have rolled threads and are assessed to demonstrate they meet the fatigue requirements in NASA-STD-5001.

(5) Fasteners manufactured from titanium alloys require additional coordination with the RFCB for approval.

Titanium alloys, such as Ti-6Al-4V (including annealed and STA conditions), cp- Ti, and other titanium alloys, have potential generic EAC or SLC failure modes that are to be addressed in the assessment with test data from flawed fasteners in the applicable service life environments.

(6) The fasteners are not made from a low fracture toughness alloy, as defined in section 3.2 in this NASA Technical Standard.

(7) Fasteners are not reworked or custom made unless the application is approved by the RFCB.

b. Include preload and its effect on flaws and cyclic stresses in the damage tolerance assessment.

c. Inspect all fracture critical fasteners by the eddy current NDE technique or use proof testing to screen for flaws.

d. Assume a flaw in the most critical location of a size consistent with NDE sensitivity or proof-test level in the damage tolerance analysis.

General NDE flaw sizes are given in NASA-STD-5009, Nondestructive Evaluation Requirements for Fracture Critical Metallic Components; but for specific guidelines on eddy current methodology, PRC-6509, Process Specification for Eddy Current Inspection, can be used as a reference.

e. Proof-load test inserts used in conjunction with fracture critical fasteners to a minimum factor of 1.2 after installation.

NASA-STD-5019A w/CHANGE 4

This would include, for example, inserts bonded or potted into composite and sandwich structures, as well as inserts installed into metallic structures. Note that composite structures require additional considerations, as given in section 7.4 in this NASA Technical Standard.

• f. Store and control fracture critical fasteners after inspection or testing to keep them isolated from other fasteners.

NASA-STD-5019A w/CHANGE 4

Tethers should be proof tested, inspected, and assessed for damage in accordance with applicable operational requirements.

Proof tests are usually performed in the operational environment, or the test levels are adjusted via an ECF. Other requirements such as NASA-STD-5001 provide proof test levels.