火箭发动机部分进气涡轮设计与流动分析

摘要

根据某型液体火箭发动机总体性能及结构要求,采用一维方法设计了部分进气、圆锥形喷嘴、单级超声速冲击式涡轮.基于求解雷诺平均的Navier-Stokes方程组,对涡轮内部流场进行全三维粘性定常仿真计算及分析,并研究了不同转子叶栅通道面积变化方式对涡轮性能的影响.结果表明:部分进气涡轮内部流动流线不规则、存在较多漩涡流动、转子叶栅激波复杂、叶片通道内分离较为严重;喷嘴通道和转子叶栅通道内总压损失均在20%以上,其中转子叶栅通道损失更大;不同转子叶栅通道面积的变化方式对涡轮总体性能影响基本不大,但收缩-扩张型通道可降低流速,缓解气流分离,对降低叶片温差应力有一定帮助.%According to the overall performance and structure requirements of a liquid rocket engine, a partial admission, conical nozzle and single-stage supersonic impulse turbine were designed with one-di-mensional engineering method. By solving Reynolds-averaged Navier-Stocks ( RANS ) equations, the three-dimensional viscous steady-state simulation and analysis are performed for the internal flow field of the turbine. For the rotor cascade channel, the influence of area varieties on turbine performance is studied. The results show that the internal streamline of the partial admission turbine is irregular, and there are more eddy flows, complicated rotor cascade shock and serious separation in the blade channel. The total pressure loss in the nozzle channel and rotor cascade channel is more than 20% , and the pressure loss in the later is greater. The variation of rotor cascade channel area has little effect on the overall performance of turbine, but the convergent-divergent type channel can reduce the flow velocity and relieve the gas flow separation, which is helpful in reducing the temperature difference stress of blade.