超声速进气道边界层吸除方案设计及实验  

Design and wind tunnel test on supersonic inlet with boundary layer bleed

作　　者：张红军[1] 忻贤钧[1] 白葵[1] 沈清[1] 

机构地区：[1]中国航天空气动力技术研究院,北京100083

出　　处：《实验流体力学》2008年第1期88-91,共4页Journal of Experiments in Fluid Mechanics

摘　　要：应用工程设计方法,结合数值模拟,设计了一种带有边界层吸除型式的超声速轴对称进气道,对进气道内流场进行了数值模拟研究,并且进行了风洞实验。研究发现,对进气道中心锥边界层进行合理流量的吸除可以明显提高进气道的总压恢复,增强了进气道的稳定工作的能力。从试验数据可知,在Ma=4.0时,进气道临界总压恢复系数达到了0.43,与不吸除比较,比常规同类进气道的临界总压恢复系数(σ=0.33)提高了约30%。通过对数值模拟结果与风洞实验结果的对比可知,二者能够基本吻合。Based on engineering design method and combined with numerical simulation, an axial symmetry supersonic inlet with boundary layer bleed was designed. The internal fluid field was investigated using numerical method and wind tunnel test was conducted. It＇ s shown that to bleed boundary layer of the center cone in a proper range of mass flow rate can improve the total pressure recovery coefficient largely and intensify stable work ability of the inlet. The total pressure recovery coefficient of the inlet reached 0.43 when Mach number reached 4.0, which had been improved by 30 % compared with the same type normal inlet without the boundary bleed. The numerical simulation result and wind tunnel test result agree well with each other.

关 键 词：边界层吸除 超声速进气道 数值模拟 风洞实验 

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] V211.753