[FCR 14] Each fracture critical part that is not of a specific hardware type as described in section 7.2 in this NASA Technical Standard and is approved as appropriate for one of the following optional approaches by the RFCB shall comply with one of the following items:
- a. Single-event fracture critical components comply with section 7.5.1 in this NASA Technical Standard.
- b. HCF components comply with section 7.5.2 in this NASA Technical Standard.
- c. Proof test approach for composite or bonded hardware complies with section 7.5.3 in this NASA Technical Standard.
- d. Fleet leading testing approach complies with section 7.5.4 in this NASA Technical Standard.
- e. Hazardous fluid containers for payloads and experiments comply with section 7.5.5 in this NASA Technical Standard.
[Rationale: Parts that comply with this requirement have had sufficient activities performed to establish adequate risk mitigation of failure caused by the presence of a flaw or crack-like defect and are approved by the RFCB.]
Use of an alternative approach requires unique rationale and approval by the RFCB as described in section 10 [FCR 26] in this NASA Technical Standard.
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7.5.1 Single-Event Fracture Critical Components
Fracture critical components with a single-event life loading history, such as pyrotechnic components, may be shown acceptable by demonstrating a factor of 1.4 on critical stress intensity factor instead of a factor of 4 on life, if all of the following conditions apply.
For single-event fracture critical components, satisfy the following items 7.5.1.a, 7.5.1.b, 7.5.1.c, and either 7.5.1.d or 7.5.1.e (as appropriate for the material and situation) to meet requirement [FCR 14] section 7.5.a in this NASA Technical Standard:
- a. The single-event loading is a single cycle or a single cycle with rapidly decaying subsequent cycles.
- b. The component is not subject to any other significant loads.
- c. The evaluation, whether by analysis or testing, and any deviations from the prescribed approaches in this section are coordinated in advance with and approved by the RFCB.
- d. Metallic components are shown by analysis to satisfy a minimum factor of 1.4 on critical stress intensity factor.
The margin is be computed as:
Margin on Critical Stress Intensity Factor = \frac{critical stress intensity factor}{(1.4 \times K_{applied})} - 1
where the:
critical stress intensity factor is usually represented as the plane strain fracture toughness, , or a parameter such as with approval of the RFCB.
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e. Both non-metallic components and metallic components satisfy the requirements of this section by using process controls that ensure the flight hardware will be represented by tests conducted on identical samples that establish the following: Tests may be used in situations where the applied loads are difficult to determine, the material properties are uncharacterized, or other factors make the damage tolerance analyses difficult.
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(1) Tests include a flaw in the worst location and orientation in the test articles.
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(2) Apply either approach A or B below to establish the components are acceptable:
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A. Use this approach when loads are known and can be readily applied to test articles.
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- i. The test load is at least 1.4 times the maximum expected flight load.
- ii. The flaw size is at least as large as the detectable sizes in NASA-STD-5009 (RTD for composite or bonded hardware, as described in section 7.4.3 in this NASA Technical Standard) for the inspection method applied to the flight hardware.
- B. Use this approach when loads are not well characterized or are difficult to apply.
- i. The flaw size is at least twice as large in all dimensions as the detected sizes in NASA-STD-5009 (twice as large as the RTD for composite or bonded hardware as described in section 7.4.3 in this NASA Technical Standard) for the inspection method applied to the flight hardwate.
- ii. The load application is to simulate worst-case flight conditions.
- iii. A sufficient number of articles are tested to ensure the test conditions approach the maximum flight conditions.
7.5.2 High-Cycle Fatigue Components
Fracture critical components operating in a potential HCF environment may be shown acceptable by establishing no HCF flaw growth. Examples of these are turbine blades, rotors, impellers, and other high-speed elements that are subject to local modes of high-frequency vibration and large numbers of loading cycles.
Satisfy the following for HCF Components to meet requirement [FCR 14] section 7.5.b in this NASA Technical Standard:
- a. Use a value for fatigue crack growth threshold that has been approved by the RFCB.
- b. Assume the initial NDE flaw size in the worst location and orientation.
- c. Propagate the flaw (by analysis or test) for 4 times the required design life using the low-cycle loads.
- d. Use the final flaw size from the calculations or test data in 7.5.2.c (above) as the initial flaw size in calculating the stress intensity factor (metallic components) or total strain (composite or bonded components) related to the HCF environment.
- (1) The metallic component is acceptable if the calculated HCF stress intensity factor is below the stress intensity factor threshold for the metallic material.
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- (2) The composite or bonded component is acceptable if the calculated net section strain (or stress) is below the no-growth threshold strain (or stress) for the composite or bonded material with RTD determined flaws.
All items 7.5.2.a through 7.5.2.d are typically performed analytically. Items 7.5.2.b and 7.5.2.c may be performed by test.
7.5.3 Proof Test Approach for Composite or Bonded Hardware
Proof test, as an optional approach, is a category available on a limited-use basis. Use of this classification should include the RFCB early in the program. The proof test classification is usually limited to payload or secondary structures. These structures should have well defined load paths, loads, and boundary conditions. The proof test should adequately load all appropriate members and sections of the structure, where necessary both in tension and compression (load reversal). In cases where shear and/or compression dominate, the proof test approach may not be appropriate because of delamination growth under these load conditions. If proof test does not adequately replicate operational conditions, this may not be an applicable approach.
Satisfy the following for the proof test approach for composite or bonded hardware to meet requirement [FCR 14] section 7.5.c in this NASA Technical Standard:
- a. Proof test the flight article to 1.2 times the limit load using one of the following:
- (1) Conduct the proof test in the appropriate environment.
- (2) Adjust the test loads using a coupon or hardware element test verified ECF.
- b. Perform pre-proof and post-proof NDE, including special visual inspection if necessary, on the hardware.
- c. Repair or replace hardware with indications of flaw growth or initiation that are discovered during proof test or with post-proof NDE.
- (1) Repeat the proof test to 1.2 times the limit load for repaired hardware.
- (2) Perform pre-proof and post-proof NDE, as well as special visual inspection if necessary, on the repaired regions.
- d. Define the threats that may cause flaws from any source that may occur to the hardware during its service life, considering all applicable flaw detection and mitigation strategies that are implemented for the flight hardware.
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- e. Develop and implement an IDMP for the hardware that assures a complete record of hardware impact or damage status and mitigates the risk of undetected damage from the threats identified in 7.5.3.d (above) for the period between post-proof NDE and launch.
- f. Establish that the largest remaining residual threat after post-proof NDE through the remainder of the service life can create damage no larger than the flaw size screened by NDE.
- g. Repeat the proof test, repair, or replace the hardware as described in 7.5.3.a through 7.5.3.c (above) if any incidents of impact or other damage occur after post-proof NDE and before launch.
- h. For re-flight hardware, repeat the proof test approach activities in items 7.5.3.a through 7.5.3.g in this NASA Technical Standard before the hardware is re-flown. Proof test loads should be limited to less than 80 percent of ultimate strength of the structure for the appropriate mode of failure, e.g., tension, compression, and shear. Structures with an ultimate safety factor of 1.5 or greater will preclude exceeding 80 percent of ultimate strength when using a test factor of 1.2. Note that the full DTA activities of section 7.4.1 in this NASA Technical Standard are not required. However, test data describing capability relative to damage or flaws will likely be necessary to assist with disposition of any flaws discovered during pre-proof NDE. Test data for capability relative to damage or flaws may also be necessary to develop NDE criteria for reportable flaws. The relevant capability is dependent on the failure mode of concern, e.g., compression-after-impact strength, delamination growth, or other. Proof tests are usually performed in the operational environment, or the test levels are adjusted via an ECF.
The composite or bonded structure should be designed so that accompanying metallic parts do not experience detrimental yielding during the proof test.
7.5.4 Fleet Leader Testing
Satisfy the following for fleet leader testing to meet requirement [FCR 14] section 7.5.d in this NASA Technical Standard:
- a. Provide the approach and rationale to the RFCB for approval before implementation.
- b. Document the approved approach in the FCP.
In cases where loading conditions are poorly defined, a ground test fleet leader program that allows use of the hardware may be feasible.
7.5.5 Hazardous Fluid Containers for Payloads and Experiments
The hazardous fluid containers category is limited to payload and experiment applications at conditions defined in requirements below. This hardware type is not part of a pressurized system nor is it intended to transfer stored fluid as part of a pressurized system.
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Satisfy the following for hazardous fluid containers for payloads and experiments to meet requirement [FCR 14] section 7.5.e in this NASA Technical Standard:
- a. The container is limited to an MDP of 152 kPa (22 psi, 1.5 atm) and a maximum volume of 0.05 m³ (1.76 ft³).
- b. An analysis is to show a positive margin against burst when a factor of 2.5 on MDP is used.
- c. Perform proof test to 1.5 MDP.
- d. Establish that no damage or detrimental deformation exists after the proof test.
- e. Establish damage tolerance against rupture and leak by satisfying sections 8 and 9 in this NASA Technical Standard for all materials, section 7.3 in this NASA Technical Standard for metallic parts, section 7.4 in this NASA Technical Standard for composite or bonded parts, and by test or analysis as approved by the RFCB for other materials.
- f. In addition to section 8 requirements in this NASA Technical Standard, perform an NDE inspection of all fusion joints in the container after proof test to determine acceptable conditions both on the surface and within the fusion joint.
- g. Perform a leak test to 1.0 times the MDP.
In instances where NDE is not feasible, the manufacturer may employ a process-control program that assures the quality of the uninspectable welds and obtain approval of the RFCB.
Proof tests are usually performed in the operational environment, or the test levels are adjusted via an ECF.
Inertial load effects (including attach points) may necessitate additional assessments beyond the items in this category.