基于导叶端弯的小展弦比燃气涡轮优化设计  

Optimization design of a low aspect ration gas turbine for curved guide vane

作　　者：毛凯[1] 李昌奂[1] 张聃 蒋建园 MAO Kai;LI Changhuan;ZHANG Dan;JIANG Jianyuan(Xi an Aerospace Propulsion Institute,Xi an 710100,China)

机构地区：[1]西安航天动力研究所

出　　处：《火箭推进》2019年第6期23-28,共6页Journal of Rocket Propulsion

基　　金：国家重大基础研究项目(613321)

摘　　要：以某型火箭发动机用亚声速小展弦比燃气涡轮为研究对象,为进一步改善涡轮内部流场,提高了涡轮效率,通过调整导叶子午端壁型线曲率、采用导叶端弯的设计方法对涡轮进行了优化设计,其作用在于减小叶片通道二次流损失,并将导叶出口压力分布进行调整,从而减小叶顶泄漏损失。基于六面体网格,采用CFX流场分析软件对优化前后结构进行了数值计算,结果表明:优化后单通道无叶顶间隙模型涡轮效率提高1.4%;采用正弯设计后,轮毂和叶顶处绝对和相对气流角显著增大,叶片中部气流角有所减小,整体分布更加均匀,消除了原型结构动叶轮毂区的流动分离;优化后全通道模型围带间隙前后压差明显降低,泄漏量从7%降低至4.75%,涡轮效率提高5.9%。In order to further improve the internal flow field and improve the turbine efficiency,a sub-sonic low aspect gas turbine for a certain rocket engine was taken as research object,the optimization was developed in this paper through the adjustment of meridional endwall’s curvature and the design of curved guide vanes.Its function is to reduce the secondary flow loss and optimize the guide outlet pressure distribution at guide vanes outlet,thereby reducing the leak loss at blade top.Base on the hexahedron grid,CFX flow field analysis software was used to calculate the turbine performance.The result shows that the optimized turbine efficiency of single channel gap-less structure increases by 1.4%;the absolute and relative airflow angles at hub and tip of the blade are significantly increased after the curved blade was used,the airflow angle in the middle of the blade is reduced,and the overall distribution is more uniform,the flow separation of the prototype structure is eliminated;the relative leakage of blade tip with maze structure gap decreases by 7%to 4.75%,turbine overall efficiency increases by 5.9%.

关 键 词：小展弦比 端弯导叶 数值计算 优化 

分 类 号：V434.21[航空宇航科学与技术—航空宇航推进理论与工程]