论文部分内容阅读
以某型航空发动机压气机2级转子叶片为例,研究了叶片的振动疲劳裂纹扩展规律。研究过程中,首先利用有限元方法分别计算了试验状态与工作状态下叶片振动导致的裂纹尖端应力强度因子范围随裂纹长度的变化;试验研究了裂纹扩展速率与裂纹长度的关系。之后,综合计算结果和试验结论,得出叶片试验状态与工作状态下的裂纹扩展规律,并与Paris公式进行了比较,发现叶片的振动疲劳裂纹扩展速率dad N是与裂纹长度a和裂尖应力强度因子范围IΔK相关的多项式,而Paris公式不能描述叶片的振动疲劳裂纹扩展现象。研究结论可进一步确定叶片的损伤容限、确定合理的叶片检修周期,为保障飞行安全奠定基础。
Taking a 2-stage rotor blade of an aero-engine compressor as an example, the law of blade’s fatigue crack growth was studied. In the course of the research, firstly, the finite element method was used to calculate the variation of the stress intensity factor range of crack tip with the crack length under the condition of the test and the working state respectively. The relationship between the crack growth rate and the crack length was experimentally studied. After the calculation and the conclusion of the experiment, the law of crack propagation under the condition of the blade and the working state is obtained and compared with the Paris formula. It is found that the dad N of the fatigue crack growth rate of the blade is the same as the crack length a and the crack tip stress The strength factor range IΔK related polynomials, while the Paris formula can not describe the blade vibration fatigue crack growth phenomenon. The conclusions of the study can further determine the damage tolerance of the blade, determine the reasonable blade maintenance cycle, and lay the foundation for the flight safety.