超声速进气道喉部附面层抽吸  

Research on boundary-layer suction in the throat of supersonic inlet

作　　者：严红明[1] 钟兢军[2] 韩吉昂[1] 冯子明[1] 于洋[2] 

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 [2]大连海事大学轮机工程学院,辽宁大连116026

出　　处：《推进技术》2009年第2期175-181,共7页Journal of Propulsion Technology

摘　　要：为研究超声速进气道喉部之后流场激波附面层干扰,采用FLUENT软件模拟了单楔角进气道在设计工况下流动情况。通过分析,提出进气道喉部抽吸。计算了三种抽吸缝大小下进气道喉部之后流场,计算结果表明,喉部抽吸能使激波稳定于喉部,通过抽吸能改善喉部之后流场状况,提高进气道性能,少量抽气不改变流场结构,加大抽气量,使喉部之后激波串转变成正激波,正激波之后流场不分离,进气道出口性能参数提高显著。To investigate shock/turbulent boundary-layer interaction after the throat of the supersonic inlet, the commercial CFD software FLUENT was exploited to simulate the flow field of the single wedge compression supersonic inlet at the design point. By analyzing the flow loss mechanism, boundary layer suction method at the throat of the supersonic inlet was introduced, and the flow fields of inlet with three kinds of suction slot dimension were simulated. The simulation results indicate that boundary layer suction in throat can fix the shock in the throat and it forms a stable flow field. It can improve the inlet performance. Less suction mass flow ratio will not change the structure of the flow field. By increasing suction mass flow, the shock train after the throat will turn to a normal shock. As there is no separation after the normal shock, the flow field is stable.

关 键 词：超音速进气道 边界层 干扰 激波 抽吸^+ 

分 类 号：V235.213[航空宇航科学与技术—航空宇航推进理论与工程]