Damage tolerant design criteria and the current use of random or flight-by-flight loading methods in the full-scale testing of aircraft structures have necessitated the development of new techniques for acquiring accurate fatigue crack growth information from fracture surface analysis. This paper describes a technique that utilized a 64-cycle constant-amplitude “marker block”, in an otherwise random load spectrum of 4469 cycles, to derive fatigue crack growth data from an aluminum alloy spar cap and two wing skin rivets. The presence of an undetected incipient crack in the spar cap, from a previous full-scale test, provided an opportunity of accounting for the entire random loading crack growth history in terms of the “marker band” spacings. It was possible to identify the marker bands in the scanning electron microscope at relatively short crack lengths of 2.6 mm for the spar cap and 0.5 mm for the rivets. A total of 89 of the 90 marker blocks applied in the test were accounted for on the spar cap fracture surface. Examination of the rivet fracture surfaces revealed 12 and 8 marker bands, respectively, over crack lengths of some 3 mm.
The circumstances surrounding the in-service failure of a cast Ni-base superalloy (Alloy 713LC) second stage turbine blade and a cast and coated Co-base superalloy (MAR-M302) first stage air-cooled vane in two turbine engines used for marine application are described. An overview of a systematic approach, analyzing the nature of degeneration and failure of the failed components, utilizing conventional metallurgical techniques, is presented. The topographical features of the turbine blade fracture surface revealed a fatigue-induced crack growth pattern, where crack initiation had taken place in the blade trailing edge. An estimate of the crack-growth rate for the stage II fatigue fracture region coupled with the metallographic results helped to identify the final mode of the turbine blade failure. A detailed metallographic and fractographic examination of the air-cooled vane revealed that coating erosion in conjunction with severe hot-corrosion was responsible for crack initiation in the leading edge area.
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