Abstract
For gas turbine engines, safe life methodology has historically been used for fatigue life management of failure critical engine components. The safe retirement limit is necessarily determined by a conservative life procedure, and many components which are discarded have a long residual life which is thereby forgone. The objective of this study is to introduce the damage tolerant design concept into the life management for aircraft engine component instead of conservative fatigue life methodology which has been used for both design and maintenance. Crack growth data were collected on a nickel base superalloy which have been subjected to combined static and cyclic loading at elevated temperatures. Stress analysis for turbine disk was carried out. It was estimated that remaining life of turbine disk component under various temperature and conditions using creep-fatigue crack growth data. As the result of life assessment, it was confirmed that retirement for cause concept was applicable for the evaluation of useful remaining life of retired turbine disk which had been designed based on conventional fatigue life methodology.
