Downloaded from https://www.everyspec.com ADS-79D-HDBK When the engine exceeds its normal operating limits, damage accumulates much more rapidly. Moreover, a rotor speed or temperature limit exceedance may require that the engine be removed and inspected. The first step in the Life Usage Monitoring processes is predicting the life of a critical part. Under normal design operating conditions, failure in turbine engine life-limited parts is caused by a combination of low cycle fatigue (LCF), high cycle fatigue (HCF), thermal fatigue, and creep. Premature failure due to external factors such as foreign object damage, corrosion, material defects, etc.
can be addressed through condition monitoring techniques described in section G.7.4. LCF results from the large stresses induced by repeated cycling of the engine between low and high power operating points. The primary engine cycle is typically startup-flight-shutdown. Smaller stress cycles within this primary cycle also contribute to LCF. For example, throttle movements at higher engine speeds cause damage through centrifugal loads excursions on rotating components. In typical turboshaft engines, LCF is the predominant cause of failure. HCF is caused by lower stresses than those associated with LCF, but at much higher frequencies. Typical causes of HCF are vibration and flutter.
29 HCF can be avoided under normal operating conditions by proper design and selection of materials. However, because military aircraft may operate under conditions conducive to foreign object damage (FOD) and inlet distortions, HCF can occur from notched blades caused by FOD and inlet distortions caused by weapons firings or rapid maneuvers. Thermal fatigue results from stresses induced by temperature gradients across components. Stress levels and frequencies are similar to those associated with LCF. Creep is the dimensional change in metal parts resulting from sustained loads at high temperatures. Creep may be avoided or controlled through proper design and selection of materials.
Determining the life of a part is predicated largely on predicting the effects of LCF and thermal fatigue. This requires conducting stress and heat transfer analyses using well-defined material properties. The expected stresses and environmental conditions for normal engine operation must also be used in these analyses. Bench, rig, and spin pit tests can then be used to validate and improve the initial analyses. Finally, developmental flight testing of the engine is used to verify the environmental conditions and stresses assumed in the analyses.
The safe life of a life-limited part is defined as the estimated amount of time before the first measurable crack appears or, in other words, there is low probability the material strength will degrade below its design ultimate value due to fatigue cracking. This estimation is based on the stresses derived from the stress analysis and the material properties (the empirical stress/cycle (S/N) data). The loads used in the stress analysis are developed based on an assumed mission profile or mix of mission profiles.
Because of the uncertainty in material properties / stresses / mission profile / desired maintenance intervals, some conservatism is used to set service life limit lower than the safe-life. With modern engine monitoring systems it is possible to determine usage much more accurately and reduce the uncertainty between the estimated life consumed (based on hour/cycles with a nominal mission profile) and the actual life consumed. The key is determining the actual number of cycles that the components experience during the mission. This can be done by a number of methods, with the 29 SAE AIR 1872A Guide to Life Usage Monitoring and Parts Management for Aircraft Gas Turbine Engines, pg. 7 119