Fatigue Lives of Specimens Representing Critical Locations in Mirage III Spars under Australian and Swiss Test Spectra

Full-scale flight-by-flight fatigue tests on Mirage III fighter aircraft wings indicated much greater differences in the flights to failure than were expected. This problem has been by conducting fatigue tests on specimens representing the major failure locations in the test wings - a blind anchor-nut hole and an interference-fit bolt hole - under flight-by-flight loading sequences corresponding to those used in the full-scale tests. Constant-amplitude fatigue tests were carried out on the blind-hole specimens to provide information for life estimation purposes, and fractographic crack propagation studies were conducted on selected blind-hole specimens. The experimentally determined lives of both types of specimen were greater under the blind hole stress spectrum. Depending on the stress scaling factor considered the life ratio obtained by grouping together the two stress spectra and the two types of specimens were between about 3 and 4.2. This compares with a ratio of 4.6 obtained for the respective full-wing scale wing fatigue tests. For the blind anchor-nut hole specimens the predicted fatigue lives were less than those obtained experimentally at high stress scaling factors, but at low stress scaling factors the converse was the case. The ratios of the predicted lives to failure were much greater than those found experimentally. Under both stress spectra, the lives of front-flange bolt hole specimens incorporating interference-fit bolts were between 3 and 4 times greater than those with clearance-fit bolts. Keywords: Australia; Fatigue(materials); Aluminum alloys; Aircraft structures; Mirage aircraft; Bolted joints; Interference-fit fasteners; Load spectra; Stress concentration; Crack propagation.