1+1/2对转涡轮用出口超音叶栅设计与试验  

DESIGN AND TEST OF A SUPERSONIC CASCADE USED FOR THE VANELESS COUNTER-ROTATING TURBINE

作　　者：季路成[1] 黄海波[1] 陈江[2] 于海力[1] 徐建中[1] 

机构地区：[1]中国科学院工程热物理研究所,北京100080 [2]北京全三维动力公司,北京100080

出　　处：《工程热物理学报》2004年第1期45-48,共4页Journal of Engineering Thermophysics

基　　金：中国科学院知识创新工程项目(KGCX2-301);国家自然科学基金(No.59906012);国家重点基础研究发展规划项目(No.G1999022305)

摘　　要：本文以1+1/2对转涡轮为背景,开展出口马赫数1.5、气流角为70°高出口马赫数涡轮叶栅设计与试验研究。研究与分析表明,尾缘厚度及尾缘附近叶表速度分布是决定上述高出口马赫数叶栅性能的关键;尾缘后约一倍叶栅出口宽度范围内,损失剧烈增加,此距离之后,总压降低趋于平缓。初步试验结果说明高出口马赫数涡轮叶栅是可行的。In order to advance the use of the vaneless counter-rotating turbine, the designing and experimental study of a high supersonic turbine cascade is presented, which its exit flow Mach number is 1.5 and its exit flow angle is 70? It is shown that the performance of the high supersonic turbine cascade is determined by the thickness of the trailing edge and the velocity distribution on the surface near the trailing edge. The loss of the cascade increases dramatically after the trailing edge in the range of one width of the throat, the loss increases lowly and smoothly out of this range. The result of the experiment done on a shock tunnel test rig illustrates that the high supersonic turbine is usable.

关 键 词：超音出口涡轮叶栅 无导叶对转涡轮 激波管风洞 

分 类 号：V235[航空宇航科学与技术—航空宇航推进理论与工程]