目的 提高航空发动机叶片抗外物损伤的性能。方法 采用薄壁件激光冲击强化工艺，对某型发动机TC4钛合金叶片包含一阶弯曲振动节线区域的表面进行处理，随后在叶片前缘一阶弯曲振动节线位置设计不同应力集中系数的缺口。参考有限元仿真软件分析结果和相关标准要求，预制应力集中系数Kt为3.2的缺口。通过力值校核和有限元仿真之间的多次迭代，明确应力测试位置与缺口危险点应力之间的关系。通过振动疲劳试验对激光冲击强化效果进行评价。通过扫描电子显微镜观察疲劳断口的形貌，采用残余应力仪对梯度残余应力进行测试，并提取相应位置的半峰全宽值，对激光冲击强化提升缺口叶片疲劳强度的原因进行分析。结果 经激光冲击强化处理后的钛合金缺口叶片在107次循环下的疲劳强度提升了63.2%；残余压应力层深度可达1.5 mm，且表层位错密度提升了67.5%；经激光冲击强化处理后钛合金缺口叶片裂纹萌生于近表面。结论 激光冲击强化引入的表层梯度残余压应力和位错增殖是缺口叶片疲劳强度提升的主要原因。

The work aims to improve the performance of aero-engine blade against foreign object damage (FOD). The surface of TC4 titanium alloy blade containing first-order bending vibration pitch line of a certain engine was processed by laser shock peening (LSP) technology of thin-walled component. The position of the blade to be processed was partitioned, and the guided wave material was pasted on the back to prevent the deformation and spallation of blade. Then, the notches with different stress concentration coefficients were designed at the first-order bending vibration pitch line of the blade leading edge. The stress gradient at the root of notch changed dramatically and the maximum stress was difficult to measure. Therefore, finite element method was used to find suitable monitoring area to characterize the stress of the notch location. According to the analysis results of finite element simulation software and the requirements of relevant standards, the notch with a stress concentration coefficient Kt of 3.2 was prefabricated. Through several iterations between stress calibration and finite element simulation, the relationship between stress test position and the stress at the notch risk point was clarified. According to the finite element simulation results, the strain gauge was pasted at the corresponding position of the blade, and the measured results further indicated that the stress at the notch could be better characterized by monitoring at other positions of the blade. The effect of laser shock peening was evaluated by vibration fatigue test. The standard required that the evaluation criterion of the blade was 107. On the premise of satisfying the cycle life, the test was carried out through step by step loading method with 106 as a cycle. The fatigue strength of titanium alloy notched blade increased by 63.2% under 107 cycles after LSP. The morphology of the fatigue fracture was observed by scanning electron microscope. The fatigue fracture of the specimen after LSP was obviously larger than that of the un-LSP specimen, and the undulating morphology was formed in the process of fatigue crack propagation. In contrast, the surface of the un-LSP specimen was relatively flat. The crack initiation of titanium alloy notched blade was near the surface after LSP. Fatigue cracks often originated from the surface of components. The surface stress state and microstructure had great effect on the fatigue performance. Therefore, the gradient residual stress was measured by residual stress meter and the value of full width at half-maximum of corresponding position was extracted. The depth of residual compressive stress layer reached 1.5 mm, and the dislocation density of surface layer increased by 67.5%. The deeper residual stress layer meant that the applied stress could be effectively balanced and the crack propagation could be delayed during the whole working process of the component. The increase of dislocation density could effectively refine the grain size, and the effect of fine grain strengthening could also improve the fatigue performance. The gradient residual compressive stress and dislocation multiplication introduced by LSP are the main reasons for the improvement of fatigue strength of notched blades.