航空发动机叶片TC4钛合金振动疲劳裂纹扩展研究及剩余寿命预测

Research on Fatigue Crack Propagation and Remain Fatigue Life Prediction of Aero-engine Blade TC4 Titaniumalloy

目的研究TC4钛合金的振动疲劳特性及寿命预测。方法通过共振疲劳试验,分析裂纹尖端应力强度因子的变化规律,计算不同应力水平下疲劳裂纹扩展的速率,建立剩余寿命预测计算模型。结果裂纹尖端的应力强度因子是表征裂纹扩展速率快慢的有效参数,与裂纹长度及应力场的大小相关。在裂纹扩展初期应力为274 MPa的条件下,裂纹扩展速率的试验值与计算值吻合较好。通过寿命预测模型计算可知,当初始裂纹为0.5 mm,最终裂纹长度达到5 mm时,在应力为274、366、422 MPa的条件下,振动循环周期分别为36 577、19 090、13 865。结论在应力比为?1的振动条件下,裂纹扩展速率随应力水平的增大而加快,同时初始裂纹长度越长,应力相同时,裂纹扩展速率提高。通过寿命预测模型,可计算出结构件的使用寿命。

Objective To study vibration fatigue performance of TC4 titanium alloy and predict its service life. Methods The resonance fatigue test was used to analyze the change regularity of stress intensity factor at the crack tip and calculate the fatigue crack propagation rate under different stress levels. The model of remain fatigue life prediction was created. Results The stress intensity factor was an effective parameter to characterize the crack propagation rate and was related to crack lengths and stresses. The calculation results of fatigue crack propagation rate at the initial were in good agreement with fatigue test results at 274 MPa. When the initial crack length was 0.5 mm and the final crack length was 5 mm, the remain fatigue life was 36 577,19 090 and stress ratio, tion will be 13 865 cycles at 274 MPa, 366 MPa and 422 MPa respectively. Conclusion Under the vibration conditions at -1 the crack propagation is accelerated with increase of stresses. If the initial crack length is longer, the crack propaga-bigger under the same stress. The remain life ofworkpiece can be calculated by model in this paper