[FCR 12] Each fracture critical metallic part that is not of a specific hardware type as described in section 7.2 in this NASA Technical Standard and is not approved by the RFCB as appropriate for an optional approach as described in section 7.5 in this NASA Technical Standard shall comply with one of the following item combinations: a and b; a and c; or a, b, and c:
- a. Develop loading spectra by complying with section 7.3.1 in this NASA Technical Standard.
- b. Perform assessment by analysis to comply with section 7.3.2 in this NASA Technical Standard.
- c. Perform assessment by test to comply with section 7.3.3 in this NASA Technical Standard.
[Rationale: Fracture critical parts need activities performed to understand the sensitivity of the part if a flaw is present. These activities can range from a direct assessment of the part’s capability with a flaw to acceptance tests that establish the part has sufficient capability to a combination of activities that provide sufficient information to mitigate the risk of failure related to undiscovered flaws.]
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. The approaches in this requirement are the preferred approaches if followed completely.
Damage tolerant assessment used as the basis for acceptance of a fracture critical metallic part establishes all of the following:
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- • The relevant critical failure mode for the part is identified.
- The appropriate load spectra are applied.
- The appropriate initial flaw size in a worst-case orientation based on the screening method implemented, in the worst location, is used.
- • Conservative material data and analysis methods are used.
- • One of the following (each of which is detailed in this section) is established:
- The part has a minimum service life factor of 4.
- The part is single loading event hardware and has a factor of 1.4 on critical stress intensity factor or residual strength.
A damage tolerance assessment is performed to understand the sensitivity of a part to flaws. The requirement is necessary to mitigate risk of failure because of flaws that may still exist after implementation of flaw screening strategies. Fatigue-crack-growth empirical data have inherent scatter. When performing damage tolerance assessments, mean values are used, not a statistical lower bound. In addition, the prediction procedures have uncertainties related to the local stress levels, stress-intensity factor calculations, load spectra, and environmental effects. Errors in local stresses and stress-intensity factor calculations are grossly magnified when crack growth rates are evaluated while using the Paris growth law. Slight misjudgments of the spectrum can lead to large effects on crack growth. To account for all of these effects, a safety factor is applied on the predicted life. Thus, the life factor of 4 provides margin on uncertainties in analysis, prediction methodologies, and material property variations. The single load event factor of 1.4 on critical stress intensity factor, fracture toughness, or residual strength provides ultimate load capability with flaws that may go undetected and is representative of the requirements in NASA-STD-5001.
7.3.1 Loading Spectra
A loading spectrum is necessary for the damage tolerance life analysis or damage tolerance life test.
Develop loading spectra according to the following to satisfy requirement [FCR 12] section 7.3.a in this NASA Technical Standard:
- a. Include all anticipated significant loadings, both cyclic and sustained, for each fracture critical part throughout its service life.
- b. Include all load levels and the number of cycles or duration during the service life of the hardware, including proof test loads.
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- c. Include the effects of the appropriate environment for each fracture critical part throughout its service life.
- d. Include the effects of preloads.
- e. Include residual stresses and any weld joint discontinuities, such as peaking and mismatch, for cyclic and sustained loads during the service life of the hardware.
- f. Include the influence of all coatings and barriers on pressure-loaded parts for any scenarios where pressure is assumed to decrease because of leakage from a crack.
- g. Include the effects of impact loads and damage from mission environments, including but not limited to credible impacts from vehicle loss of external surface mass, MMOD, EVA inadvertent contacts, and EVA tool impacts during assessments of external structures and components.
Include the worst-case allowed or weld joint peaking and mismatch effects for damage tolerance assessments by analysis or test. The assessment analysis or test is to capture the effect of peaking and mismatch on stress gradients affecting crack growth and fracture. Standard tensile strength tests of ductile materials are not adequate to assess these conditions.
Proof load factors are listed in NASA-STD-5001 and may exist in program-specific requirements. Proof tests are usually performed in the operational environment, or the test levels are adjusted via an ECF.
7.3.2 Assessment by Analysis
Satisfy the following to perform assessment by analysis to meet requirement [FCR 12] section 7.3.b in this NASA Technical Standard:
- a. Assume that the initial flaw that could be present and undetected in the part is the size and shape that is not screened by NDE, proof test, or process control and is in the worst location and orientation.
- b. Use analysis methods and computer programs that are approved by the RFCB, e.g., NASGRO®, for predicting flaw growth, life, and critical flaw sizes.
Note that when the available analysis ability to simulate crack growth is invalid, assessment by test (section 7.3.3 in this NASA Technical Standard) is required.
- c. 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.
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- (1) If the loading sequence of high/low loads is unknown, then damage tolerance analysis is to show that the stress intensity at limit load is less than the critical stress intensity factor or residual strength at the end of 4 lifetimes.
- (2) 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:
- A. Reserve capability against fracture by meeting either a lower bound critical stress intensity factor or residual strength at the end of 4 lifetimes.
- B. 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 because of SLC or EAC during the service life are addressed in item 7.3.2.f below.
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d. Use flaw growth rates that are greater than or equal to the average values without implementing retardation effects on flaw growth rates in the damage tolerance analysis.
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e. Use critical stress intensity factor and cyclic threshold stress intensity range (ΔKth) values that are less than or equal to the average values.
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f. For metallic alloys susceptible to EAC or SLC or both, satisfy all of the following:
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(1) Use the lower bound value of stress intensity factor threshold for assessment of EAC (K_EAC or K_IEAC as appropriate) and SLC if the material exhibits these behaviors in the application conditions.
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(2) 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 (1) above at the end of 4 lifetimes. Requirement 7.3.2.f is intended to preclude susceptible metallic alloy flight hardware from experiencing time-dependent, i.e. da/dt, crack growth.
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g. If NASGRO® is used:
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(1) Bk is either set to zero, or Bk is set such that Kc at the part thickness is less than or equal to the Klc value.
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(2) Values of Bk resulting in Kc > KIc require further understanding of the constraint condition for the crack situation and may be used with approval of the Technical Authority or RFCB.
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- h. Use fracture properties subject to all of the following:
- (1) From sources or testing that are approved by the RFQB.
- (2) Representative of the material process condition.
- (3) Representative of weakest material orientation in the part (unless material orientation is fully traceable throughout the design and service life).
- i. If material data needed for the damage tolerance assessment are not available, one of the following is to be accomplished:
- (1) Obtain the data by material testing.
- (2) If the source of the data to be used is from the literature, conduct an assessment to show that conservative results are obtained using that available data.
Section 8.1 in this NASA Technical Standard specifies flaw screening methods. The damage tolerance assessment is to address flaws that are not screened by the screening method applied to the flight hardware.
The NASGRO® computer program is an approved analysis tool for the damage tolerance life assessment of metallic spaceflight hardware. Other computer programs or analysis methods are acceptable with prior approval by the RFCB. The NASGRO® material database contains fracture mechanics properties for several materials that can be used with concurrence from the RFCB.
Standard NASA damage tolerance analyses are deterministic, and experience has shown these deterministic methods to be adequate. The probabilistic method uses knowledge (or assumptions) of the statistical variability of the damage tolerance variables to select criteria for achieving an overall success confidence level. Any proposed use of probabilistic damage tolerance analysis or criteria to meet fracture control requirements is considered an alternative approach as described in section 10 in this NASA Technical Standard and is approved by the RFCB on a case-by-case basis.
7.3.3 Assessment by Test
Perform assessment by test according to the following to satisfy requirement [FCR 12] section 7.3.c 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.
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- c. Perform the test(s) with initial flaws in the worst location and orientation.
- d. Establish by testing that the components survive 4 lifetimes, including section 7.3.2.c.(1) and 7.3.2.c.(2) requirements in this NASA Technical Standard, without failure (leak or fracture instability).
Testing may be supplemented by analyses that, in conjunction or augmented by test correction factors, assess all applicable effects causing increased crack growth.
- e. Test in conditions that account for the service environments.
- f. A sufficient number of tests is performed to establish a representative result considering variability of material damage tolerance data.
The approved approach is to be documented in the FCP. Formal documentation in the FCP facilitates in-depth technical review and approval. Testing of coupons and pre-flawed structural elements representative of the flight hardwate damage tolerance condition may be an acceptable approach to establish damage tolerance for metallic fracture critical parts. Together, the testing and any supplemental analyses are to establish that equivalent section 7.3.2 requirements in this NASA Technical Standard are met.