涡轮叶栅前缘气膜冷却对气动参数影响的数值研究  

Numerical simulation of influence of leading edge film-cooling on aerodynamic parameters in turbine cascade

作　　者：颜培刚[1] 王松涛[1] 韩万金[1] 

机构地区：[1]哈尔滨工业大学能源科学与工程学院,黑龙江哈尔滨150001

出　　处：《推进技术》2004年第1期44-47,共4页Journal of Propulsion Technology

基　　金：国家"九七三"基金资助项目 (G1 9990 2 2 3 0 7)

摘　　要：采用具有TVD性质的三阶精度有限差分格式、自由型曲面技术以及多区域网格算法,对前缘带有三排冷气孔的涡轮导叶进行了气膜冷却数值计算,分析了前缘冷气喷射对涡轮气动参数的影响,描述了叶型表面冷气射流的运动规律。结果表明,冷气喷射导致了流量、马赫数和温度较为明显的变化,前缘和吸力面获得的绝热效率高于压力面,压力面冷气射流的运动规律比较复杂。型面压力只在冷气孔区域有明显的波动。Finite difference scheme with three-order precision and TVD property, arbitrarily-curved surface technique and multi-block computational method were employed. The film-cooling flow field in turbine cascade with three rows of coolant holes at leading edge was calculated and its influence on aerodynamic parameters was discussed under two blowing ratios. The coolant ejection near blade surface was illustrated to show the difference of cooling film formation on the stagnation region, pressure and suction surfaces of the leading edge. The result shows that inlet mass flux and local mach number are both reduced appreciably as coolant flux increases. Profile pressure is not changed obviously in the whole, with the exception in the local region near the coolant hole with a pressure variation, which is greater on the suction side. The adiabatic effectiveness on the stagnation line and suction surface are higher than that on pressure surface. The discipline of coolant flow on the pressure side of leading edge is very complicated.

关 键 词：涡轮叶栅 薄膜冷却 气动力参数 有限差分理论 数值仿真 

分 类 号：V430[航空宇航科学与技术—航空宇航推进理论与工程]