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

DSTO-TR-14770.120.120.10.10.080.080.060.060.040.040.020.0200 0.02 0.04 0.06 0.08 0.1Type I EPS (measured flaw depth) mm.0.1200 0.02 0.04 0.06 0.08 0.1 0.12Type II EPS mm.99.9999.90.1990.080.060.040.0200 0.02 0.04 0.06 0.08 0.1Type I EPS (measured flaw depth) mm.959080705030201051.1Type I EPS (measured flaw depth) mm.Type II EPS mm.Type IV EPS mm..010.001 0.01 0.1 1EPS flaw depth mm.Figure 33Type II EPS plotted against the measured initial flaw depth (Type I EPS). Type VI EPS plottedagainst Type I EPS and the Type IV EPS plotted against Type II EPS. The line in each graphindicates the one-to-one ratio. Finally the cumulative probability plot for the three types of EPSis shown with the best fit curves.An examination of the mean and SD log data given in Table 10 along with the plotted data in thefinal graph of Figure 33 indicates that although the Type II EPS values are on average smaller thanthe Type IV values, the lower Type IV SD log values results in the small value at the 1 in 1000 EPSsize, which is not all that different to the value predicted by the measurement of the flaw depth. Itmay be that with sufficient data, for the purposes of prediction, the use of the flaw depth in thesepeened specimens, and possibly peened F/A-18 structure, may be sufficient for the prediction of aparts life when considering the 1 in 1000 life point.The size of the 1 in 1000 EPSs, although an ‘imaginary’ measure and not an actual crack size,appears to be about the depth where the peening ceases to be fully effective and on these groundsit may be predicted that there will be instances where complex shaped aircraft components whichhave been ‘repaired’ by peening will not receive the expected life extension, and therefore be a32