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Appendix A: Synopses of Air Force Aging Aircraft Structural Histories
Pages 85-110

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From page 87...
... Although the 707 had a design 87 life goal of 20,000 flights or 60,000 flight hours, the Air Force did not specify a design service life goal for the KC-135. In 1962 the Air Force decided to perform a full-scale fatigue test of the KC-135 to quantify its expected life.
From page 88...
... Also, the panel expressed concern about the long-term effectiveness of the cold-worked fastener holes in the 7178-T6 aluminum lower surfaces of the outer wing (i.e., this stemmed from the fact that some small fatigue cracks had been reported in some outer wing fastener holes in one specific aircraft that had been inspected) and recommended several actions, including assessing the need to replace these lower wing surfaces and the station 733 joint closure rib that has had a lot of problems with corrosion and SCC.
From page 89...
... The only alternative to grounding (or replacing the lower wing surface) is to protect the structural safety through frequent, careful, and very burdensome inspections of all highly-stressed fastener holes in the spanwise splices to detect and repair cracks before they reach critical size.
From page 90...
... . Thus, the IRT determined safety limits for the C-5 wings based on slow crack growth from an assumed maximum probable initial manufacturing damage in a spanwise splice fastener hole and also by estimating the onset of WFD based on the fatigue test results.
From page 91...
... By the time of the January 1975 review by the Division Advisory Group committee, the operational spectrum for the C-5A had been changed from the original 14-mission spectrum to a new spectrum called the "representative mission profiles," or the RMP spectrum, so as to more nearly reflect the actual and planned aircraft use. Also, the tear-down inspection of the full-scale fatigue test wings had revealed more spar web cracking than had previously been thought to exist, and the damage tolerance analyses had been refined to include the effects of shear loading on crack growth rates.
From page 92...
... The design life goal of this modified wing and body structure was 12,000 flight hours, and it was fatigue tested to 72,000 cyclic test hours or six lifetimes during the 1960s. After the test, the tear-down inspection revealed about 222 cracks.
From page 93...
... 1 . 1 1~1 God - 1 t ~l l ~ 1 ~1 1 1 ~ 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 I Flight hours in thousands I 2020 2030 2040
From page 94...
... Although the original ground fatigue tests and ground vibration tests showed no problems and the natural frequencies of the tail were beyond those at which acoustically driven problems would be expected, flight experience showed that highcycle fatigue cracking occurred very early in the service life. After considerable effort on the part of the contractor, the problem was found to be caused by the fact that the production tails had gaps between the skins and the titanium substructure.
From page 95...
... However, the increased use severity will increase the inspection burden, and some of the wing inspections could become particularly onerous because of the current lack of NDE capability to inspect for small cracks without removal of the fasteners. McDonnell Douglas is currently using the results of the tear-down inspection of the fatigue test aircraft and crack growth analyses to obtain a better estimate of the actual service life expectancy of the F-15.
From page 96...
... When the full-scale development began in January 1975, the airframe was designed to the ASIP requirements in MIL STD-1530A and the damage tolerance requirements in MILA-83444. Materials and stress levels were selected to preclude the need for structural inspections throughout the AGING OF U.S.
From page 97...
... A primary purpose of this new test was to identify the areas of the structure that had become fatigue critical as a result of the more severe and different use of the aircraft. The test revealed about a dozen new critical areas by October 1989, when it had reached 7,330 cyclic test hours.
From page 98...
... The six problem areas that were highlighted were - cracking of the vertical tall attachment bulkhead at fuselage station 479 - cracking of fuel vent holes of the lower wing skin - cracking of the wing attach bulkhead at fuselage station 341 cracking of the upper wing skin fastener problems on the horizontal tall support boxbeam · cracking of the ventral fin In each case, the current repair and replacement concepts were descnbed. Concerns about future fatigue cracking were expressed, along with the possibility of hidden corrosion.
From page 99...
... Although the goal of the original full-scale fatigue test was to achieve two lifetimes (i.e., 12,000 hours) of the original design spectrum (which it did obtain with some problems)
From page 100...
... During the November 1996 review with this committee, the representative of the A-10 system program director listed the current known fatigue cracking problems on the service aircraft. In addition to some of the service problems listed above, he noted that fatigue cracking was occurring in the forward fuselage upper crown skin and at the lower wing skin fastener holes and pylon stud holes at wing station 23.
From page 101...
... The original design life goal for the commercial 707 aircraft was 20,000 flights and 60,000 flight hours. During the refurbishment and modification of the initial commercial 707s to the E-8 configuration, corrosion was found to be quite extensive in the aircraft fuselage.
From page 102...
... Selected fastener holes were examined further using stereoscopic microscopes. No fatigue cracks were found in any of the fastener holes or other structural details.
From page 103...
... During the 1970s it was discovered that the damage tolerance of the outer wings was severely degraded due to fatigue cracking plus some faulty depot maintenance actions by some commercial contractors. In the 1979 to 1981 time
From page 104...
... are -3 g to +7.33 g, and the original design life goals were 4,000 flight hours and ten years of service. At the time of the original design, the Air Force had not yet developed and implemented damage tolerance design requirements.
From page 105...
... / IFS 770 bulkhead and ~ ~ engine bay frame / FS 496 Nacelle former Wing carry-through box was completed successfully. A full-scale fatigue test program was initiated in 1968, with the entire program lasting about six years.
From page 106...
... Over the years, fatigue cracks have been detected and were subsequently repaired prior to proof testing in 25 different areas of the wing carry-through structure, the wing pivot fitting, the horizontal stabilizer support structure, the fuselage station 496 nacelle former, the main landing gear support structure, and in the fuselage structure. Nevertheless, proof test failures have not been totally avoided.
From page 107...
... Fatigue was not recognized as a serious concern by the Air Force until the fatigue failure of the front spar of the wing of an operational aircraft in 1968. This resulted in a modification to the aircraft and the subsequent fatigue testing of two full-scale wing and carry-through structures during 1968 to 1969.
From page 108...
... In 1970 there was a failure of this lower wing skin on an F-5 stationed at Williams AFB in Chandler, Arizona. It was caused by a fatigue crack, which originated at the trailing-edge radius at wing station 26 and grew to a critical size of about 0.20 in.
From page 109...
... Since the DADTA and the resulting recovery program in the late 1970s, the lower wing skins have been replaced with thicker skins made from 7075-T73 aluminum, the fastener holes and drain holes in high-stressed areas have been cold worked, and additional full-scale fatigue testing has been performed. Also, the T-38s have been replaced by F-16s in the Thunderbird demonstration team, and the Air Force no longer uses the T-38 in DACT use and has replaced LIF use with IFF use.
From page 110...
... material change/redesign force-wide modification of upper longerons force-wide inspection/repair of lower longerons · Fuselage forgings bulkheads at fuselage stations 325 and 362; formers at fuselage stations 332, 487, and 508 temporary repairs · Horizontal stabilizer core corrosion attributable to water intrusion - improved bonding being implemented · Landing gear strut door - core corrosion attributable to water intrusion - superplastic formed/diffusion-bonded titanium redesign current preferred spare Based on the above, it is apparent that the San Antonio ALC will continue to face a major challenge to protect the safety and prolong the service life of the T-38 for another 25+ years as indicated in Chapter 2 of this report. Most of the cracking problems noted are safety-of-flight concerns.


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