超声速/高超声速双拐点喷管设计  

Contour design of super/hypersonic dual-inflection nozzle

作　　者：郭善广[1] 王振国[1] 赵玉新[1] 柳军[1] 

机构地区：[1]国防科学技术大学高超声速冲压发动机技术重点实验室,长沙410073

出　　处：《航空动力学报》2012年第12期2742-2748,共7页Journal of Aerospace Power

基　　金：国家自然科学基金(11072264)

摘　　要：为实现直连式试验台、高温风洞等试验设备的多马赫数运行,提出了双拐点喷管设计方法.喷管分2段设计,第1段共用,采用3次B-Spline函数描述喷管轴线马赫数分布.首先采用特征线方法求解Eul-er方程,得到无黏的理想喷管型面.其次采用参考温度方法求解边界层位移厚度,对无黏壁面进行修正得到实际壁面.共用段喷管出口的平行均匀流作为第2段喷管设计的初值.为验证设计方法的可行性,设计了中间马赫数为3.0,出口马赫数分别为4.0,4.5和5.0的双拐点喷管,并采用雷诺平均的Navier-Stokes方程对设计的喷管流场进行数值模拟.计算结果表明:喷管出口流场均匀,试验菱形区的马赫数误差小于1.2%.该方法提高了喷管设计精度,保证消波干净,为直连式试验台、高温风洞等设备的多个喷管共用一套动力系统提供了基础.In order to achieve multi-Mach number run for the direct-connected test bench and high temperature wind tunnel,a straightforward technique has been developed to quickly determine a design of the super/hypersonic dual-inflection nozzle.The nozzle was divided into two sections to be designed,and the first was shared for different test Mach number.B-Spline functions were used to describe the Mach number distribution along the axial line of the nozzle.The preliminary nozzle contour was generated by a new computer program that solved the axisymmetric Euler equations using the method of characteristics.The final nozzle design was obtained by adding the boundary-layer thickness which was solved with the reference temperature method.An example of a dual-inflection supersonic nozzle design was employed to illustrate the technique with a computational fluid dynamics calculation.The simulation results indicate that desired Mach numbers are obtained at the nozzle exit,and the good flow quality is attained for different nozzles within Mach number error less than 1.2% in the test rhombus region.The present technique improves the design precision of the converging-diverging nozzle,cancels waves completely,and achieves the nozzles with different exiting Mach numbers which shares subsonic section and a portion of supersonic section.

关 键 词：喷管设计 超声速 双拐点 型面 特征线 消波 

分 类 号：V231.1[航空宇航科学与技术—航空宇航推进理论与工程]