亚声速进气道出口流场畸变控制研究  

Investigation on flow distortion control of a subsonic inlet

作　　者：王健[1] 李应红[1] 张百灵[1] 

机构地区：[1]空军工程大学工程学院,陕西西安710038

出　　处：《推进技术》2010年第2期143-146,共4页Journal of Propulsion Technology

基　　金：空军重点科研项目;空军工程大学工程学院优秀博士学位论文创新基金项目(BC08001)

摘　　要：一种亚声速进气道出口流场存在较严重的总压畸变,为改善其出口流场品质,抑制总压畸变,首先分析了引起总压畸变的原因,即进气道扩张段内边界层发生分离;其次提出了在进气道内安装叶片式涡流发生器的流动控制方法,并进行了仿真验证;最后进行了进气道缩比模型的风洞试验。试验结果表明,在进气道设计马赫数(0.65)和非设计马赫数(0.21)条件下,安装叶片式涡流发生器后,在流量系数0.4～0.85范围内,进气道出口流场的综合畸变指数分别平均降低14.7%和23.8%,因此验证了流动控制方法的有效性。In order to suppress the total pressure distortion at the exit plane of a subsonic inlet,the cause was analyzed and it indicated that the boundary layer separated severely in the inlet diffuser section. Then the flow control method of vane-type vortex generators （VG） was proposed,and the simulation results showed that vane-type VG were effective to suppress flow separation. Finally,the wind tunnel experiment of inlet scale model with or without vane-type VG was developed. When operating at design Mach number （0.65） and off-design Mach number （0.21）,respectively,the vane-type VG averagely decreased the inlet distortion index by 14.7% and 23.8%,respectively with flow coefficient of 0.5-0.85. This demonstrated that the proposed flow control method is effective.

关 键 词：涡轮喷气发动机 进气道 进气道畸变 边界层分离 风洞试验 

分 类 号：V211.48[航空宇航科学与技术—航空宇航推进理论与工程]