北京航空航天大学学报 ›› 2020, Vol. 46 ›› Issue (10): 1890-1898.doi: 10.13700/j.bh.1001-5965.2019.0471

• 论文 • 上一篇    下一篇

燕尾榫连接结构微动疲劳全寿命预测方法

徐可宁1, 李雯1, 黄勇2, 余庆陶3, 马国佳3, 胡文颖1   

  1. 1. 中国航空发动机研究院, 北京 101304;
    2. 北京航空航天大学 航空科学与工程学院, 北京 100083;
    3. 中国航空制造技术研究院, 北京 100024
  • 收稿日期:2019-09-02 发布日期:2020-10-29
  • 通讯作者: 李雯 E-mail:mosquato@buaa.edu.cn
  • 作者简介:徐可宁 女,博士,高级工程师。主要研究方向:航空发动机结构强度振动可靠性;李雯 女,博士,研究员。主要研究方向:机械传动与摩擦学、微动损伤多尺度建模与仿真技术。
  • 基金资助:
    国家自然科学基金(51705490,5187600);装备预研重点实验室基金(614290802081706)

A fretting fatigue total life prediction method for dovetail attachment

XU Kening1, LI Wen1, HUANG Yong2, YU Qingtao3, MA Guojia3, HU Wenying1   

  1. 1. Aero Engine Academy of China, Beijing 101304, China;
    2. School of Aeronautic Science and Engineering, Beihang University, Beijing 100083, China;
    3. AVIC Manufacturing Technology Institute, Beijing 100024, China
  • Received:2019-09-02 Published:2020-10-29
  • Supported by:
    National Natural Science Foundation of China (51705490,5187600); Pre-research Key Laboratory Fund for Equipment (614290802081706)

摘要: 微动损伤使航空发动机榫连接结构疲劳寿命显著降低。以钛合金Ti-6Al-4V燕尾榫连接结构为例,提出一种适用于复杂结构微动疲劳全寿命预测方法。基于修正的Manson-McKnight方法和多轴疲劳理论,疲劳损伤参数由等效应力参数(ESP)表征,微动疲劳裂纹萌生位置和成核寿命通过有限元分析(FEA)和ESP预测。基于断裂力学理论和最大周向应力准则,提出微动疲劳裂纹扩展数值模拟方法,建立微动疲劳扩展寿命与裂纹长度函数关系,依据裂纹终值长度预测微动疲劳扩展寿命。结果显示:钛合金Ti-6Al-4V燕尾榫连接结构微动疲劳裂纹扩展角预测值与实验值均为18°,裂纹生长方向预测值与实验值相符;微动疲劳全寿命(成核寿命+扩展寿命)预测值在实验值的2倍分散带内;最大拉伸载荷对榫连接结构的微动疲劳全寿命影响显著,在相同应力比下,最大拉伸载荷从18 kN变化到24 kN,钛合金Ti-6Al-4V燕尾榫连接结构微动疲劳全寿命降低1个数量级。

关键词: 航空发动机, 燕尾榫连接结构, 微动损伤, 微动疲劳, 寿命预测

Abstract: Fatigue life of aero-engine dovetail attachment can be significantly reduced by fretting damage. Taking Ti-6Al-4V alloy aero-engine blade dovetail attachment as an example, a fretting fatigue total life prediction method for complex structure is proposed. A fatigue damage parameter was defined as an Equivalent Stress Parameter (ESP) based on modified Manson-McKnight method and multiaxial fatigue theory. Crack initiation position and nucleation life were evaluated by ESP and multiaxial stresses obtained from Finite Element Analysis (FEA). A numerical simulation method of fretting fatigue crack growth is proposed based on linear elastic fracture mechanics and maximum hoop stress criterion. From the simulated crack growth results, the function relationship between fretting fatigue propagation life and crack length was established, and the fretting fatigue propagation life was determined by the crack length at failure. The results show that the crack growth trajectory predicted by the simulation correlates well with that in a tested Ti-6Al-4V dovetail component-both have a crack kink angle of 18°. The estimated fretting fatigue total life (nucleation + propagation) of a dovetail under different fretting conditions by using the proposed numerical method matches well with test results, as the predicted total life is within 2 times of error range. Maximum tensile load has significant influence on crack nucleation and propagation life, and under the same stress ratio, the fretting fatigue total life of the Ti-6Al-4V dovetail attachment reduces one order of magnitude as the maximum tensile load increases from 18 kN to 24 kN.

Key words: aero-engine, dovetail attachment, fretting damage, fretting fatigue, life prediction

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