Section 3.3.0. Proof Test Determinations
Tiffany and Masters [1965] first suggested that the
conventional structural proof test could be used to inspect for crack damage
that would eventually lead to catastrophic failure. These techniques were first applied to rocket motor cases and
tankage as a result of numerous missile launch failure at Cape Canaveral. Air Force acceptance of this proof test
philosophy has been stimulated by the inability of alternate non-destructive
inspection tools to reliably detect cracks of near-critical size. The Air Force in the recent past has
employed the proof test as a means of determining the maximum possible initial
flaw that could exist in the structural subsystems identified in Table 3.3.1.
Note that almost all of the examples cited represent the application of
the proof stress techniques as an In-service Inspection. Buntin [1971], Cowie [1975], Horsley, et al.
[1976], Gunderson [1974] and Albrectsen & Aitken-Case [1976]
document the Table 3.3.1 and other Air Force uses of
the crack-inspection proof test. White,
et al. [1979] documents the recent Navy proof test of an A-7 arresting
hook; this proof test is periodically repeated to ensure the continuing
structural integrity of the component.
The proof test concept for all applications has been to size or
eliminate the life degrading damage so that the structure would maintain its
required level of structural integrity throughout a defined period of usage.
However, due to substantially different technical requirements, the proof
testing techniques employed in each case were different. The technical
requirements that establish the type of tests conducted have been structural
geometry, material properties, type of crack damage present in the structure,
as well as the crack growth mechanism.
Table 3.3.1. Proof Testing of Aircraft
Structures
System
|
Subsystem
|
Damage
|
Special
Techniques
|
F-111
|
Lower surface of inner wings and pivot fittings
|
Potential forging defects propagated due to fatigue in D6AC
steel
|
Upwing bending at -40° F after every 1,000
hours of flight
|
B-1A
|
F-101 (Development) engine combustor case
|
Pores and inclusion stringers in circumferential butt welds
in Inconel 901 alloy
|
Internal pressure to 200% operating pressure
|
B-52D
|
Center and inner wing structure
|
Fatigue and stress corrosion cracks nucleated during southeast Asia service in 7075-T6 and
7079-T6 aluminum alloy structure
|
Down and up-wing bending at ambient temperature
|
C-141
|
Main Landing gear (cylinder)
|
Hydrogen entrapped during refurbishment
|
500 hours of continuous static loading to initiate and
propagate cracks to failure
|
A-7
|
Carrier arresting hook (Navy)
|
Fatigue cracking initiated during service
|
Repeat periodically
|