Abstract

The life of turbine blades often determines the life of the aero-engine, and creep rupture is one of the main fracture forms of turbine blades in high temperature environments. In order to accurately predict the life of turbine blades, creep tensile tests of smooth specimens were carried out according to typical loads of 760 °C/800 MPa, 900 °C/445 MPa, 980 °C/400 MPa, 1000 °C/283 MPa and 1050 °C/210 MPa. With the help of crystal plasticity theory, creep constitutive equations with temperature interpolation were constructed. The FSI heat transfer analysis was performed to obtain the creep life of the turbine blades,. The temperature difference on the blade had reached 300 °C. With creep constitutive model and temperature boundary, the creep analysis of the blade was carried out and the creep process was revealed. The dangerous section of the global-model was obtained, and the sub-model was used to re-analyze the dangerous area. This paper reveals the deformation of the blade after creep failure: the middle part of the blade shrinks which is also the potential damage area, the top part expands, and the trailing edge of the blade tip produces a radial displacement of 0.313 mm after creep failure. The creep life of the turbine blades is determined to be 91 h. Finally, a method to determine the creep life of turbine blades was given based on skeletal point stress method.

摘要

涡轮叶片的寿命往往决定着航空发动机的寿命，而蠕变断裂是涡轮叶片在高温环境下的主要断裂形式之一。为了准确预测涡轮叶片的寿命，按照760℃/800MPa、900℃/445MPa、980℃/400MPa、1000℃/ 283 MPa 和 1050 °C/210 MPa。借助晶体塑性理论，建立了带温度插值的蠕变本构方程。执行 FSI 传热分析以获得涡轮叶片的蠕变寿命。刀刃上的温差已经达到了三百摄氏度。利用蠕变本构模型和温度边界，对叶片进行蠕变分析，揭示蠕变过程。得到global-model的危险部分，使用子模型重新分析危险区域。本文揭示了蠕变失效后叶片的变形：叶片中部收缩也是潜在的损伤区域，顶部膨胀，蠕变失效后叶尖后缘产生0.313 mm的径向位移. 涡轮叶片的蠕变寿命确定为 91 小时。最后，给出了一种基于骨架点应力法的涡轮叶片蠕变寿命测定方法。蠕变破坏后叶尖后缘产生0.313 mm的径向位移。涡轮叶片的蠕变寿命确定为 91 小时。最后，给出了一种基于骨架点应力法的涡轮叶片蠕变寿命测定方法。蠕变破坏后叶尖后缘产生0.313 mm的径向位移。涡轮叶片的蠕变寿命确定为 91 小时。最后，给出了一种基于骨架点应力法的涡轮叶片蠕变寿命测定方法。