涡轮叶栅双排孔气膜冷却数值模拟  

NUMERICAL SIMULATION OF FILM-COOLING IN TURBINE CASCADE WITH TWO ROWS OF STAGGERED COOLANT HOLES

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学院,黑龙江哈尔滨150001

出　　处：《工程热物理学报》2004年第5期757-760,共4页Journal of Engineering Thermophysics

基　　金：国家科科技部973项目资助(No.G1999022307)

摘　　要：采用具有三阶精度TVD性质的有限差分格式、自由型曲面技术以及分区网格算法,对某型具有冷气孔形状的涡轮叶栅进行了全三维N-S方程数值求解,描述了相邻两排冷气射流在叶栅吸力面形成的冷却气膜以及壁面附近冷却射流运动的特点,分析了不同吹风比和喷射角度情况下冷却绝热效率的分布规律。结果表明,在较大的吹风比和喷射角下,交错排列的两排冷却射流运动规律非常复杂,在两排孔之间的区域与冷气孔下游区域冷却气膜的形成规律具有明显的区别。A finite difference scheme with third order precision and TVD property and a technique of multi-block grid are applied in this paper. Simulation of film-cooling on the suction surface of a turbine cascade with two staggered rows of coolant holes is performed based on the three-dimensional NS solver. The characteristics of film-cooling of the coolant from the two rows of coolant holes on blade suction surface are displayed as well as the coolant trajectory along downstream. The distribution discipline of film-cooling adiabatic effectiveness is analyzed under different blowing ratios and ejection angles. The result shows that the transportation of two rows of staggered coolants is complicated especially under large blowing ratio and angle, and the formation of film-cooling adiabatic effectiveness in the region between two rows of holes is distinctively different from that in the downstream of the second row.

关 键 词：涡轮 气膜冷却 分区网格 冷气射流 绝热效率 

分 类 号：V231.3[航空宇航科学与技术—航空宇航推进理论与工程]