定几何混压式轴对称超声速进气道气动特性数值仿真和实验验证  

Numerical Simulation and Experimental Validation of Flow in Mixed-Compression Axisymmetric Supersonic Inlet with Fixed-Geometry

作　　者：谢旅荣[1] 郭荣伟[1] 

机构地区：[1]南京航空航天大学能源与动力学院,江苏南京210016

出　　处：《航空学报》2007年第1期78-83,共6页Acta Aeronautica et Astronautica Sinica

基　　金：国家863高技术项目(2003AA723020)

摘　　要：针对一种马赫数为4一级的定几何混压式超声速轴对称进气道进行了数值仿真研究,并和风洞试验结果进行对照,验证了本文所采用计算方法的可靠性。利用CFD方法获得了进气道激波系分布、内通道流场分布和沿程静压分布,并对Ma=4下稳定亚临界状态进行了分析。研究结果表明:①超临界状态下,随着进气道出口反压的提高,结尾激波系向喉道方向移动,结尾激波损失减小,总压恢复系数提高;②迎角的增加对进气道的迎风侧和背风侧影响增大,结尾激波系由对称分布向一边倾斜的趋势增大,背风侧的承受反压能力下降,总压恢复系数随之下降;③随着来流马赫数的增加,激波损失加大,总压恢复系数随之下降,同时由于激波角变小,激波也越靠近外唇罩,溢流减小,流量系数增大,在激波贴口后流量系数基本保持不变;④通道内的静压分布曲线清晰地反映了内通道沿程激波系情况;⑤在大于贴口马赫数工作时,结尾激波系被推出唇口的情况下,由于滑流层作用出现一个类似外压缩式的气动通道,从而存在稳定的亚临界状态。In order to get a good knowledge of the flow in a mixed-compression axisymmetric supersonic inlet with fixed geometry at Ma= 2.5 to 4.0, a numerical study is done in the paper. The investigation focuses on the structure of the shocks and the characteristics of the internal flow, and the distribution of static pressures along the duct is also given. Compared with the experimental data, the reliabilities of the CFD in this paper are verified. Results indicate： （1） At the supercritical operation, the augment of backpressure at the exit of the inlet induces terminal shocks to move towards the throat. As a result, the intensity of terminal shocks weakens and the total pressure recovery coefficient at the exit of the inlet increases. （2） The structure of terminal shocks changes from symmetric to asymmetric with the increase of angle of attack so that the backpressure ratio at the exit of the inlet descends gradually and the total pressure recovery coefficient decreases. （3） With the rise of the Mach number of free stream, the total pressure recovery coefficient declines and the mass ratio ascends gradually and maintains 1.0 when the oblique shock attaches the cowl lip. （4） The flow pattern can be analyzed by the distribution of static pressure coefficients along the duct. （5） When the Mach number of free stream exceeds the Mach number of design point, at the subcritical condition, a bow shock stands ahead of the cowl lip, and the flow pattern of the inlet is similar to that of the external compression inlet so that the inlet can still operate stably.

关 键 词：航空航天推进系统 轴对称进气道 定几何混压式进气道 数值仿真 内流场 

分 类 号：V211.3[航空宇航科学与技术—航空宇航推进理论与工程]