Fatigue Crack Growth in 7050T7451 Aluminium Alloy Thick Section ...

DSTO-TR-14771. IntroductionDSTO research into material factors likely to be of importance to the F/A-18 fatigue life havehighlighted the critical role played by surface condition in determining the service fatigue life ofaircraft structure. Previous coupon test programs have addressed the differences betweenpolished, ‘as-machined’ and peened surfaces in terms of life when loaded by several differentwing root bending moment spectra; Molent et al , 2000; Sharp & Clark, 2001. This report examinesthe application of glass bead peening to aluminium alloy 7050T7451 thick section plate typical ofareas in the F/A-18 which have been treated to extend their fatigue life. Work similar to this haspreviously been carried out and reported in Sharp & Clark, 2001 for some wing root bendingspectra. The specimens reported on in this paper were loaded with more representative wing rootbending spectrum derived from the F/A-18 FT55 centre fuselage fatigue test being carried out inCanada, Simpson, 2002. Whereas the previous testing compared lives at a single peak spectrumstress, this examination reports on several peak stresses in order to establish the trend in the LifeImprovement Factor (LIF). In addition, a more detailed examination of the crack growthmorphology including detailed quantitative fractography on the largest cracks, at these differentpeak spectrum stresses is presented. The use of several peak stress levels produced life dataapplicable in the estimation of the LIFs for the most critical structure in the F/A-18 aircraft. TheFT55 spectrum used had additional marker loads added to aid quantitative fractography.This project was developed out of a desire to improve the understanding of fatigue crackingencountered in service and test components from the F/A-18 aircraft. To simulate the peenedsurface of the AL7050T7451 material in the aircraft, the surfaces of the test coupons were preparedto simulate those of 7050 components fitted to the F/A-18.2.1 Introduction2. BackgroundPeening treatments are widely used in mechanical and aeronautical engineering to improve thefatigue lives of components. The process is normally associated with harder metals such as steels,nickel and titanium alloys. Although this is generally the case, simply configured aluminium alloycomponents have for many years been peened, for example aluminium alloy propeller blade hubs.The application of this method to complex aluminium alloy components such as those found inthe F/A-18 is however a relatively recent occurrence. This has resulted in reports of extensivevariation in the fatigue life results for these high strength aluminium alloy peened componentsand, in some cases (Clayton & Clark, 1988), a decrease in fatigue life has been observed. Suchvariability naturally raises concern that the peening process developed for use with hard materialssuch as steels might not always be suitable for peening high-strength aluminium alloys which aremuch softer.An extensive array of literature dating back to early in the twentieth century examining the effectsof peening on steels and other hard alloys exists. An excellent review of this literature may befound in the book produced by the Metal Improvement Company, Inc of New Jersey, USAentitled “Shot Peening Applications” eighth edition, 2001. In a general sense most of this isrelevant although the softness of the aluminium alloys compared to these materials results insome special problems which need to be addressed by an improved understanding of the wayfatigue cracking propagates in a peened aluminium alloy.1