MFN Trainer Column
in Vol. 5 - May Issue - Year 2004
Peening in the Age of Damage Tolerance
Shlomo D. Ramati
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by Shlomo D. Ramati of IAI, Israel Aircraft Industries
The advent of “Damage Tolerance” (DT) has raised questions within aerospace as to the economic desirability of saturation peening of aircraft parts.
Damage Tolerance is defined as “the ability of the structure to sustain damage in the form of cracks without catastrophic consequences, until such time that the damaged component can be repaired” (Commercial Aircraft Requirement, D. Broke).
The DT method is based on the assumption that micro-cracks exist in all parts but their size is infinitesimally smaller than the reliably detectable crack in standard NDT (non destructive testing) methods.
The military specification and consequently the American FAA (FAR regulations) as well as the European JAA have adopted this method when justifying a structure for certification.
The assumed starting crack size is 1 mm. This assumption results in a heavier conservative design that does not take into consideration crack initiation time. Since the depth of the crack is beyond the compressive layer that results from saturation peening, peening cannot be taken into account when calculating the fatigue life of the part.
The life is calculated by growing the assumed crack to a critical size. This is done using data that has been established for crack growth da/dn versus stress intensity factor k in different environments. The data is fed into a computer with the expected flight induced stress spectrum and the growth program run. These models include gusts and retardation due to cold working at the crack tip after a gust load.
The first major inspection is at half this lifetime. When the NDT is done and if cracks are found then one compares them to the expected crack lengths and based on that the next inspection time is determined. This half life inspection interval is an important characteristic of the design and a major selling point when one can extend it.
Some companies calculate the fatigue life using the older traditional methods of fatigue ratings in addition to the above. These calculations take into account surface finish, finishing processes and peening. In addition landing gears and some actuators are still fatigue rated this way.
It makes sense, for example, when there are many special processes involved in a landing gear. It is well known that Hard Chrome plating induces residual tensile stresses in the base material as well as having micro-cracks (chicken wire or mud cracks) inherent to the process in the chrome layer itself. In the traditional methods these deleterious effects are taken into account by multiplying the surface co-efficient by 0.9 or 0.85. On the other hand shot peening creates a compressive residual stress and therefore slows the crack initiation and growth thereby having a multiple greater than 1 normally 1.1 or 1.15 the effects are arithmetic so that by peening before chrome plating we insure optimal fatigue life.
In summation it should be obvious that though one does not get credit for the peening when certifying the A/C in the DT era, after fatigue testing the advantages of the peening become evident when the cracks if any are far smaller than the predicted DT model.
Finally when cracks do appear the parts will have to be repaired or changed as no pilot will agree to fly with cracks in structural members. The peening will pay off in maintenance costs and integrity reputation of the A/C.
Author: Shlomo D. Ramati of IAI