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    Fatigue Analysis of Cold-Worked and Interference Fit Fastener Holes

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    Crack initiation and crack growth behavior were determined experimentally for aluminum and titanium specimens with fastener holes that were either cold worked or were propped by interference fit fasteners. The specimens were subjected to a randomized flight-by-flight spectrum. Analytical procedures were evaluated, based on correlation with the test data. These procedures included elastic-plastic analysis which was utilized to determine the stress-strain distributions surrounding the fastener holes. Estimates of elastic proportional limits in tension and compression were based on material cyclic stress-strain characteristics. Purely elastic analysis was used to determine the KT and stress gradient for the propped fastener holes with various ratios of plate-to-fastener modulus of elasticity. Finite-element model elastic-plastic computer results provided the stress-strain distributions for the cases of superimposed cold working and external loading, and superimposed interference fit fasteners and external loading. Approximate analytic equations were developed to fit the finite-element model computations. These equations were used to calculate the stress-strain excursions that would occur during flight-by-flight spectrum fatigue loading. The stress-strain excursions were then used to enter strain-life curves to compute crack initiation. They were also used to determine stress intensities to enter da / dn versus ΔK curves to compute crack growth. The delay in crack initiation due to beneficial compressive residual stresses that would be induced during spectrum loading was accounted for by using a stress ratio correction factor based on strain-life data generated at different stress ratios. Crack growth retardation due to periodically applied high loads was accounted for, using the Wheeler plastic zone model. The comparison of the crack initiation life calculations to the test data was favorable if initiation was defined as the development of a 0.25-mm crack. Good agreement between the crack growth calculations and the test data was also obtained for growth from a 0.25-mm crack.


    stresses, strains, aluminum alloys, crack propagation, fatigue (materials), cyclic loads, aircraft, tensile properties, mathematical prediction, stress analysis, residual stress, plastic deformation, titanium alloys

    Author Information:

    Rich, DL
    Lead engineer and branch chief, Technology, McDonnell Aircraft Company, McDonnell Douglas Corporation, St. Louis, Mo

    Impellizzeri, LF
    Lead engineer and branch chief, Technology, McDonnell Aircraft Company, McDonnell Douglas Corporation, St. Louis, Mo

    Committee/Subcommittee: E08.05

    DOI: 10.1520/STP27993S