湍流普朗特数在高超声速绕流中的修正  

Correction of Turbulent Prandtl Number over Hypersonic Flows

作　　者：刘景源[1] LIU Jingyuan(Aircraft Engineering College,Nanchang Hangkong University,Nanchang 330063,Jiangxi,China)

机构地区：[1]南昌航空大学飞行器工程学院,江西南昌330063

出　　处：《弹箭与制导学报》2024年第1期1-5,共5页Journal of Projectiles,Rockets,Missiles and Guidance

基　　金：国家自然科学基金项目(12362025);气动院高超声速气动力/热技术重点实验室基金项目(GT202206236)资助。

摘　　要：为精确模拟高超声速激波/湍流边界层干扰导致的复杂非平衡湍流流动,提出一种湍流普朗特数修正模型。应用数值模拟及理论分析方法,对高超声速来流马赫数为9.22的平板、压缩拐角等绕流进行数值分析,评估了所提出的湍流普朗特数修正模型。数值模拟结果与实验数据及Kays湍流普朗特数模型的对比表明:对高超声速复杂流动,湍流普朗特数应进行修正,提出的经湍流非平衡参数修正的湍流普朗特数修正模型与原模型及Kays模型相比,给出的壁面压强、壁面热流更精确,壁面最大热流相对误差小于6%。To investigate complex non-equilibrium hypersonic shock wave/turbulent boundary layer interaction flows,a turbulent Prandtl number correction model was proposed.The proposed turbulent Prandtl number correction model was evaluated by numerical simulation and theoretical analysis methods to the flows around hypersonic flat plate,compression corners at freestream Mach number 9.22.The comparison of numerical simulation results with experimental data and the Kays’s turbulent Prandtl number model indicates that for hypersonic complex flows,the turbulent Prandtl number should be corrected.The proposed turbulent Prandtl number correction model,which is modified by turbulent non-equilibrium parameters,demonstrates more accurate skin pressure and heat flux compared to the original model and Kays’s model,the maximum wall heat flux relative error is less than 6%.

关 键 词：高超声速 激波/湍流边界层干扰 气动热 湍流普朗特数 数值模拟 

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]