火箭推进

随着可重复使用发动机的研制,必须对涡轮叶片的高周疲劳问题开展深入研究。提出了一种针对局部进气冲击式涡轮在气流激振力作用下的动态响应理论计算模型,将叶片简化为悬臂梁结构,同时考虑离心力及阻尼的影响,建立其受迫振动微分方程及求解方法,同时提出了一种局部进气涡轮叶片气流激振力的简化模型。以某发动机涡轮叶片为研究对象,分别采用理论模型及三维流固耦合仿真的方法对叶片进行动态响应分析,通过两种方法的结果对比来验证理论模型的正确性。结果表明,与三维仿真结果相比,采用理论模型方法计算得到的叶片根部平均动应力低3.39%,响应频率高1.62%,响应最大幅值低11.4%。从而验证了理论模型的准确性,可以采用该方法开展火箭发动机涡轮叶片高周疲劳设计。

With the development of reusable rocket engines, the high cycle fatigue of turbine blades must be studied in depth. A theoretical model to calculate the dynamic response of partial admission turbine blades under the excitation force was proposed in this paper, and the vibration equation and solution method were established. Both theoretical method and fluid-solid coupling three-dimension simulation method were applied to an engine turbine blade. Compared to three-dimension simulation method, the average dynamic stress at blade root calculated by theoretical method is 3.39% lower, the response frequency is 1.62% higher and the maximum response amplitude is 11.4% lower. Thus, the accuracy of theoretical method is verified, and this method can be applied to the high cycle fatigue design of rocket turbine blades.