航空发动机压气机叶片振动疲劳裂纹扩展规律研究  

Vibration fatigue crack propagation law of aero-engine compressor blade

作　　者：李春旺[1] 武晓亮[1] 柴桥[1] 张忠平[1] 

机构地区：[1]空军工程大学理学院,西安710051

出　　处：《应用力学学报》2016年第3期384-388,542,共5页Chinese Journal of Applied Mechanics

基　　金：国家自然科学基金(51575524);陕西省自然科学基金(2015JM5240);空军工程大学理学院博士后创新基金(2013BSKYQD09)

摘　　要：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。An aero-engine compressor blade is taken as example, vibration fatigue crack propagation law is studied. In the studying, finite element method is used to investigate that the stress intensity factor range at crack tip changes with crack length, both for test state and for working state. Relationship between crack progress rate and crack length is obtained through tests. By synthesizing the calculated results and the tested data, the crack propagation law for both test and working state is obtained. Furthermore, the difference between the crack propagation law and Paris formula is compared. It is found that the vibration fatigue crack propagation law of the blade is a polynomial, Paris formula cannot use to describe the vibration fatigue crack propagation. The conclusions of the present paper can be a base for determining damage tolerance and repair period and for ensuring flight safety.

关 键 词：含裂纹叶片 应力强度因子 有限元方法 裂纹扩展 损伤容限 

分 类 号：V235.1[航空宇航科学与技术—航空宇航推进理论与工程]