局部进气冲击式压力级涡轮设计优化研究  

Design optimization research on impulsive pressure level turbine with partial admission

作　　者：李瑜 朱东华[1] 许开富[1] 付瑜[1] LI Yu;ZHU Donghua;XU Kaifu;FU Yu(Xi＇ an Aerospace Propulsion Institute,Xi＇an 710100,China)

机构地区：[1]西安航天动力研究所,陕西西安710100

出　　处：《火箭推进》2018年第4期16-22,共7页Journal of Rocket Propulsion

基　　金：国家863项目(2015AA7053026)

摘　　要：依据液体火箭发动机涡轮泵原理,建立了两级局部进气冲击式压力级涡轮的设计方法。该方法可以根据涡轮进出口边界条件、转速和结构尺寸等参数,完成涡轮的一维设计,并输出叶型的几何数据和流动性能参数,再结合三维数值模拟进行验证。按照涡轮总体设计要求,完成了某小流量高压比涡轮的原始设计,根据三维数值模拟的结果,对原始设计的涡轮叶型进行了优化,涡轮效率提高了2%。在全周结构上进行了三维数值模拟验证,优化后的两级局部进气冲击式压力级涡轮满足涡轮总体设计要求。A method for designing an impulsive pressure level turbine with two-stage partial admission was established based on the turbine-pump principle of the liquid rocket engine. The one dimensional design of the turbine can be completed according to the boundary condition of the air inlet and outlet,revolutions per minute,and the structure size of the turbine,and then the geometry data of the turbine blade profile and the flow parameter are exported. Finally three-dimensional numerical simulation is employed to validate the design result. The proto type of a turbine with low mass flow rate and high pressure ratio was designed in accordance with the total design desire of the turbine. On the base of the three-dimensional numerical simulation results,the turbine blade profiles of the original design were optimized.The efficiency of the turbine was increased by 2%. The three-dimensional numerical simulation of whole circumferential model was performed for the flow. The result shows that the optimized impulsive pressure level turbine with two-stage partial admission can satisfy the total design requirements of the turbine.

关 键 词：局部进气 冲击式压力级 涡轮 优化 

分 类 号：V434.211[航空宇航科学与技术—航空宇航推进理论与工程]