A long-term test run followed by room-temperature fatigue test was performed for the first-stage high-pressure turbine blades of an aero-engine. The investigation showed the appearance of cracks at the trailing edge of the blade coated with AlSiY. The failure mechanism was investigated by macroscopic observation, fracture analysis, metallographic analysis, hardness testing and structural evolution analysis by electron backscatter diffraction (EBSD), kernel average misorientation (KAM) and strain analysis. The results showed that thermal fatigue cracks initiated and propagated on the trailing edge during the test run and cracks further propagated during the fatigue test. Brittle and hard phases including M23C6 carbides, μ phase and γ'(Ni₃Al) phase precipitated in the inter diffusion zone with a high dislocation density and maximum hardness and strain, resulting in thermal fatigue cracks initiation during frequent start-stop of the aero-engine test run. The cracks propagated outward along the β-NiAl grain boundaries with interconnected γ'(Ni₃Al) phases in coating, and propagated inward along the lathlike precipitates including M23C6 carbides, M6C carbides and μ phase in the secondary reaction zone to the grain boundaries of the substrate. Chainlike M6C carbides locally precipitated on the substrate grain boundary, causing grain boundary embrittlement and further promoting cracks propagation.

对航空发动机的第一级高压涡轮叶片进行了长期测试，然后进行了室温疲劳测试。研究表明，在涂有AlSiY的叶片后缘出现了裂纹。通过宏观观察，断裂分析，金相分析，硬度测试和结构演变分析（通过电子背散射衍射（EBSD），核平均取向错误（KAM）和应变分析）研究了失效机理。结果表明，热疲劳裂纹在试验运行期间在后缘开始并扩展，而裂纹在疲劳测试期间进一步扩展。脆性和硬质相，包括M23C6碳化物，μ相和γ'（Ni ₃Al）相以高位错密度和最大的硬度和应变沉积在相互扩散区​​内，导致在航空发动机试验运行频繁起停期间引发热疲劳裂纹。裂纹在涂层中沿具有相互连接的γ'（Ni ₃ Al）相的β-NiAl晶界向外扩展，并在第二反应区中沿包括M23C6碳化物，M6C碳化物和μ相的板条状沉淀物向内扩展至Cd的晶界。基板。链状M6C碳化物局部沉淀在基体晶界上，引起晶界脆化并进一步促进了裂纹的扩展。