壁温对压缩拐角流动影响的数值模拟研究

压缩拐角是高超声速飞行器上的典型非连续区域,其分离/再附结构对局部热环境有较大影响。本文采用基于Navier-Stokes方程的自研程序,开展了典型压缩拐角外形的气动热环境数值模拟研究,分析了不同壁温条件下压缩拐角的热环境分布规律。计算结果表明:随着壁温升高,流场整体结构变化不大,但由于近壁面流体物性参数变化较大,导致拐角分离涡分离点向前移动、再附点向后移动,拐角处干扰区扩大;压缩拐角大部分区域的热流随着壁温的升高而减小,且干扰区内热流减小的幅度比无干扰区更大,但热流并不完全遵循随壁温升高而减小的规律。另外,通过对比变壁温计算热流和热壁修正公式修正热流,发现热壁修正公式在干扰区存在精度降低、适用性不足的问题。本文的研究进一步加深了壁温对压缩拐角流动影响的认识,缩小了热壁修正公式的适用范围。

Numerical simulation of wall temperature effect on compressive corner flow

Compression corner is a typical discontinuous region on hypersonic vehicle,and its separation/reattachment structure has a great influence on the local thermal environment.In this paper,a self-developed program in-house DNS code is used to simulate the aerodynamic thermal dynamics of a typical compression corner in different wall temperatures.The results show that:with the increase of the wall temperature,the overall structure of the flow field changes little,but due to the large changes of the physical parameters of the fluid near the wall,the separation point of the corner separation vortex moves forward and the reattachment point moves backward,and the interference area at the corner expands;the heat flow in most regions of the compression corner decreases with the increase of the wall temperature,and the decrease range of the heat flow in the interference area is greater than that in the non interference area.However,the heat flux does not completely follow the law of decreasing with the increase of wall temperature.In addition,by comparing the calculation of heat flux with variable wall temperature and the correction of heat flux with hot wall correction formula,it is found that the accuracy of hot wall correction formula is reduced and the applicability is insufficient in the interference area.The research in this paper further deepens the understanding of the influence of wall temperature on the corner flow,and reduces the application scope of the hot wall correction formula.