Damage-tolerance evaluation has been interpreted in the past as a means to allow continued safe operation in the presence of known cracking. This interpretation is incorrect. No regulations allow the strength of the structure to be knowingly degraded below ultimate strength(1.5 x limit). The damage-tolerance evaluation is merely a means of providing an inspection program for a structure that is not expected to crack under normal circumstances, but may crack in service due to inadvertent circumstances. If cracks are found in primary structure, they must be repaired. The only allowable exception is through an engineering evaluation, which must show that the strength of the structure will never be degraded below ultimate strength operations or in-service conditions After many major fatigue failures in the 1950s on both military and commercial aircraft, the most notable of which were the DeHavilland Comet failures in early 1954, the U.S. Air Force (USAF) initiated the Aircraft Structural Integrity Program(ASIP)in 1958(Ref 10). The fatigue methodology adopted in the AsIP was the reliability approach, which became known as the"safe-life method This safe-life approach, used in the development of USaF aircraft in the 1960s involved analysis and testing to four times the anticipated service life. On the commercial scene, another philosophy, " fail safety, was introduced in the early 1960s, and a choice between safe -life and fail-safe methods was allowed by commercial airworthiness requirements. However, it was found that the safe-life method did not prevent fatigue cracking o thin the service life, even though the aircraft were tested to four lifetimes to support one service life(i.e, scatter factor of 4 ) One notable example is the F-lll aircraft 94, which crashed in 1969(Ref 11). The F-11l aircraft had a safe- life of 4000 flight hours. However, a material defect caused the F-lll aircraft, which used high-strength steel (ultimate tensile strength of 1655 to 1793 MPa, or 240 to 260 ksi, toughness of about 66 MPa Vm, or 60 ksi vin )for the wing box(Fig 4). The defect was not observed during inspection, and a fatigue crack initiated and grew for only about 0.38 mm(0.015 in. ) The aircraft was flown for 107 flights safely, at which time catastrophic failure occurred, causing the destruction of the aircraft stayed with wing stayed with airplane Wing pivot fitting Area of anomaly (b 0.905in. Brittle fracture 0.236 in deer 0.282 in thick 0.015 in. fatigue crack growth Material anomalyDamage-tolerance evaluation has been interpreted in the past as a means to allow continued safe operation in the presence of known cracking. This interpretation is incorrect. No regulations allow the strength of the structure to be knowingly degraded below ultimate strength (1.5 × limit). The damage-tolerance evaluation is merely a means of providing an inspection program for a structure that is not expected to crack under normal circumstances, but may crack in service due to inadvertent circumstances. If cracks are found in primary structure, they must be repaired. The only allowable exception is through an engineering evaluation, which must show that the strength of the structure will never be degraded below ultimate strength operations or in-service conditions. After many major fatigue failures in the 1950s on both military and commercial aircraft, the most notable of which were the DeHavilland Comet failures in early 1954, the U.S. Air Force (USAF) initiated the Aircraft Structural Integrity Program (ASIP) in 1958 (Ref 10). The fatigue methodology adopted in the ASIP was the reliability approach, which became known as the “safe-life” method. This safe-life approach, used in the development of USAF aircraft in the 1960s, involved analysis and testing to four times the anticipated service life. On the commercial scene, another philosophy, “fail safety,” was introduced in the early 1960s, and a choice between safe-life and fail-safe methods was allowed by commercial airworthiness requirements. However, it was found that the safe-life method did not prevent fatigue cracking within the service life, even though the aircraft were tested to four lifetimes to support one service life (i.e., scatter factor of 4). One notable example is the F-111 aircraft 94, which crashed in 1969 (Ref 11). The F-111 aircraft had a safe-life of 4000 flight hours. However, a material defect caused the F-111 aircraft, which used high-strength steel (ultimate tensile strength of 1655 to 1793 MPa, or 240 to 260 ksi, toughness of about 66 MPa m , or 60 ksi in ) for the wing box (Fig. 4). The defect was not observed during inspection, and a fatigue crack initiated and grew for only about 0.38 mm (0.015 in.). The aircraft was flown for 107 flights safely, at which time catastrophic failure occurred, causing the destruction of the aircraft