Damage tolerance concept is widely used in aerospace industry for designing structural components against fatigue. In this study, an attempt is made to predict fatigue crack growth behavior in an aero-engine material under a standard spectrum load sequence using a closure independent crack driving force parameter, K* and compare with experimental results. Initially, constant amplitude (CA) fatigue crack growth rate (FCGR) behavior of GTM720, a nickel based super alloy was determined at stress ratios, R = σ_min/σ_max ranging from R = 0.1 to 0.7. Then, empirical fatigue crack growth law was derived from this CA-FCGR data in terms of two parameter crack driving force, K*. Further, the fatigue crack growth behavior was predicted through cycle-by cycle approach in a pre-cracked compact tension (CT) specimen of GTM720 alloy under standard cold-TURBISTAN spectrum load sequence using K* as the crack driving force. Also, experimental fatigue crack growth behavior under the same load sequence was determined in a CT specimen and compared with predicted results. A fairly good correlation was observed with predicted and experimental results. Predicted fatigue crack growth life was conservative and the fatigue crack growth life ratio, N_pred/N_expt was about 0.92.

损伤容忍概念在航空航天工业中广泛用于设计抗疲劳的结构部件。在这项研究中，我们尝试使用封闭无关的裂纹驱动力参数K *来预测航空发动机材料在标准频谱载荷序列下的疲劳裂纹扩展行为，并将其与实验结果进行比较。最初，GTM720的恒定幅度（CA）疲劳裂纹扩展速率（FCGR）的行为，基于镍的超合金在应力比确定的，R =σ_分钟/σ_最大R = 0.1至0.7。然后，根据该CA-FCGR数据，根据两个参数裂纹驱动力K *得出经验疲劳裂纹扩展定律。此外，在KTM作为裂纹驱动力的情况下，在标准的冷TURBISTAN频谱载荷序列下，通过逐周期循环方法对GTM720合金的预裂纹致密拉伸（CT）试样中的疲劳裂纹扩展行为进行了预测。同样，在CT样品中确定了在相同载荷序列下的实验性疲劳裂纹扩展行为，并将其与预测结果进行了比较。观察到的预测结果和实验结果都具有很好的相关性。预测的疲劳裂纹扩展寿命是保守的，疲劳裂纹扩展寿命比N _pred / N _expt约为0.92。