激波诱导轴对称气动矢量喷管壁面静压分布的试验  

Experimental investigation of wall static pressure distributions of shock induced axisymmetric fluidic vectoring nozzle

作　　者：金捷[1] 雷金春[1] 廖华琳[2] 赵景芸[2] 刘志刚[2] 

机构地区：[1]北京航空航天大学航空发动机数值仿真中心,北京100083 [2]中国燃气涡轮研究院,成都610500

出　　处：《航空动力学报》2007年第10期1700-1703,共4页Journal of Aerospace Power

基　　金：航空科学基金(01C24005)

摘　　要：在落压比3～10,次流相对流量0%～20%的情况下,对扩张段开缝的激波诱导轴对称气动矢量喷管模型试验件的壁面静压分布进行了试验.结果表明：次流的注入形成了周向壁面静压差,使喷管内流场呈现明显的三维流动特征;落压比和次流相对流量对喷管的壁面静压分布有较大的影响,并在一定条件下影响到收敛段的流动;在落压比4,次流相对流量大于15%时,激波有可能会碰到周向距次流注入最远处的壁面,使其气流分离并引起壁面静压的升高.　A scaled experimental investigation has been conducted to determine the wall static pressure distributions of a shock induced axisymmetric fluidic vectoring nozzle with a slot in its divergent section when the nozzle pressure ratio（NPR） was 3～10 and the flow flux ratio of the secondary flow injection to the primary flow（Ws/Wp） was 0%～20%.The results indicate that the internal flow characteristics are complex 3-Dimensional,for the secondary flow injection resulting in discrepancy of the circumference wall static pressure.NPR and Ws/Wp influence greatly the wall static distributions and under certain conditions they would influence the flow of the convergent section of the nozzle.When NPR=4,Ws/Wp≥15%,the shock waves may hit the lower flap which is farthest to the secondary flow injection in circumference direction,separating flows and bringing on higher wall static pressure.

关 键 词：航空、航天推进系统 气动矢量喷管 激波诱导 壁面静压分布 模型试验 

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]