在某航空发动机大修分解检查时发现，其Ⅱ级涡轮导向器结合环出现了严重变形及开裂现象，采用宏观和微观形貌观察、断口分析、化学成分分析、硬度测试等方法对其变形及开裂原因进行分析。结果表明：结合环安装孔内表面裂纹为高周疲劳裂纹；导向叶片轴颈和结合环安装孔的配合间隙及导向叶片下缘板间的间隙偏小，且导向叶片和结合环的线膨胀系数不同，使得在高温工况下轴颈不能在安装孔内自由移动，在热应力作用下，结合环向中心挤压而发生变形并形成疲劳裂纹源；由温度变化引起的间隙周期性变化是导致疲劳裂纹扩展的主要原因；适当减小Ⅱ级涡轮导向叶片轴颈尺寸，以增大其与结合环安装孔的配合间隙并适当增大导向叶片下缘板间的间隙，可避免此类故障的产生。

During the overhaul decomposition inspection of an aircraft engine, it was found that serious deformation and cracking occurred in the coupling ring of stage Ⅱ turbine guide. The cause of the deformation and cracking was analyzed by macro- and micro-morphology observation, fracture analysis, chemical composition analysis and hardness test. The results show that the cracks in the inner wall of the installation hole of the coupling ring were high cycle fatigue cracks. The fit clearance between axle neck of the guide vane and installation hole of the coupling ring, and the gap between lower edge plates of the guide vane were both small; the linear expansion coefficients of the guide vane and the coupling ring were different and therefore the axle neck could not move freely in the installation hole under high temperature condition. Under the thermal stress, the coupling ring was extruded to the center, resulting in the deformation and fatigue cracking. The periodic change of the clearances caused by temperature variation was the main reason for fatigue crack growth. The similar accidents could be effectively avoided by properly reducing the size of the axial neck of stage Ⅱ turbine guide vane to enlarge fit clearance between the axle neck and the binding ring installation hole as well as the gap between lower edge plates of the guide vane.