The Effect of Peening on the Fatigue Life of 7050 Aluminium Alloy

DSTO-RR-0208
420MPa 390MPa 330MPa
Depth <strong>of</strong> Material to be Removed
(mm)
10.00
1.00
0.10
0.01
0 2000 4000 6000 8000 10000
Simulated Flight Hours (SFH)
Figure 33: A typical plot showing how much material needs to be removed from an un-peened region
to ensure a 1/1000 chance <strong>of</strong> a crack being present. <strong>The</strong> curve is a combination <strong>of</strong> 3σ
fatigue starter and mean crack growth slopes for each stress.
A similar plot for the peened specimens has a higher 3σ initial flaw size but slower fatigue
crack growth rates (ie lower slopes).
<strong>The</strong>se plots form the basis for the repair procedure as they indicate how much material must
be removed from the surface, at a given stress, before the peening process should be
performed. To take full account <strong>of</strong> the peening LIF any “damaged surface material” must be
removed. This ensures that no cracks are present in the surface which would reduce the
effect <strong>of</strong> the repair peening. <strong>The</strong> LIF varies with stress, but at the stresses representing this
region (420MPa) a minimum LIF=1.5 can be expected.
Based on the amount <strong>of</strong> material that could be removed (definite limits based on FE analysis)
and the NDI limits an individual aircraft can be lifed as a safe-life aircraft or a safety-byinspection
aircraft. <strong>The</strong> approach is shown in Figure 34.
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