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AFGROW | DTD Handbook

Handbook for Damage Tolerant Design

  • DTDHandbook
    • About
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    • Sections
      • 1. Introduction
        • 0. Introduction
        • 1. Historical Perspective on Structural Integrity in the USAF
        • 2. Overview of MIL-HDBK-1530 ASIP Guidance
        • 3. Summary of Damage Tolerance Design Guidelines
          • 0. Summary of Damage Tolerance Design Guidelines
          • 1. Summary of Guidelines
          • 2. Design Category
          • 3. Inspection Categories and Inspection Intervals
          • 4. Initial Damage Assumptions
            • 0. Initial Damage Assumptions
            • 1. Intact Structure Primary Damage Assumption
            • 2. Continuing Damage
            • 3. Fastener Policy
            • 4. In-Service Inspection Damage Assumptions
            • 5. Demonstration of Initial Flaw Sizes Smaller Than Those Specified
          • 5. Residual Strength Guidelines
          • 6. Required Periods Of Safe Damage Growth
          • 7. Illustrative Example Of Guidelines
        • 4. Sustainment/Aging Aircraft
        • 5. References
      • 2. Fundamentals of Damage Tolerance
      • 3. Damage Size Characterizations
      • 4. Residual Strength
      • 5. Analysis Of Damage Growth
      • 6. Examples of Damage Tolerant Analyses
      • 7. Damage Tolerance Testing
      • 8. Force Management and Sustainment Engineering
      • 9. Structural Repairs
      • 10. Guidelines for Damage Tolerance Design and Fracture Control Planning
      • 11. Summary of Stress Intensity Factor Information
    • Examples

Section In-Service Inspection Damage Assumptions

The basic rationale used to write assumed sizes following an in-service inspection is essentially the same as for the case of intact structure.  Once it is established that reliance on in-service inspection is required to ensure safety, the damage size assumed to exist after an in-service inspection is that associated with the appropriate level of NDI capability, as opposed to that associated with initial manufacturing inspection capability.  In special cases where specific part removal at the depot is economically warranted, the contractor may recommend that this action be taken.  In this case, the assumed damage subsequent to part removal and inspection may be smaller than that associated with in-service inspection capabilities.  It may in fact be the same as in the original design, providing the same inspection procedures as used in production are used and certified inspection personnel perform the inspection.

Figure 1.3.6 and Table 1.3.3 summarize the in-service post inspection damage sizes as a function of conditions and thickness, from JSSG-2006 Table XXXII.  With fasteners installed and sufficient accessibility to the location, the maximum undetectable damage size is 0.25 inch of uncovered length at fastener holes.  Depending upon part thickness, this damage may be a through or part-through flaw.  The flaw size was established based on limited available inspection reliability data where the inspection was performed on the assembled aircraft as opposed to the part level inspection performed during production fabrication.  These assumptions are considered to be applicable for penetrant, magnetic particle, and ultrasonics.  Because of lack of sensitivity, X-ray is not considered appropriate for detecting tight fatigue cracks and thus is not applicable to these flaw size assumptions.


Figure 1.3.6.  Summary of Initial-Flaw Sizes for Structure Qualified as In-Service-Inspectable

Table 1.3.3.  In-Service Inspection Initial Flaw Assumptions


Inspection Method

Initial Flaw Assumption

Off-aircraft or on-aircraft with fastener removal

Same as initial

Same as initial

On-aircraft without fastener removal

Penetrant, magnetic particle, ultrasonic, eddy current

For t £ 0.25”, 0.25” through thickness flaw at holes;
For t £ 0.25”, 0.50” through thickness flaw at other locations;
For t > 0.25”, 0.25” radial corner flaw at holes; 
For t > 0.25”, 0.25” deep x 0.50” long surface flaw at other locations

On-aircraft with restricted accessibility


For slow crack growth, non-inspectable
For fail-safe structure, primary load path failed


At locations other than holes or cutouts, a flaw size of surface length 0.50 inch is assumed to be representative of depot level capability.  Where visual inspection is performed on the assembled aircraft, the minimum assumed damage is an open through the thickness crack having an uncovered length of 2 inches.  This value was established based on visual inspection reliability data derived from inspection of large transport type aircraft during fatigue testing and subsequent teardown inspection, shown in Figure 1.3.7.

Figure 1.3.7.  Development of Minimum NDI Detection for Visual Inspection