超声速球锥型飞行器气动热数值计算与研究

发布时间：2018-06-23 21:10

本文选题：超声速 + 球锥体　；参考：《南京理工大学》2015年硕士论文

【摘要】：本文基于边界层理论,采用经典的理论经验公式和数值传热相耦合的方法,对零攻角超声速球锥型飞行器气动热进行模拟计算,得到其表面热流密度值及飞行器内部温度值分布,较为精确地分析飞行器的气动热环境,为超声速飞行器外形设计优化和头部气动热设计、热防护提供了重要的理论参考。一方面对超声速圆锥绕流流场进行求解,确定球锥外形飞行器超声速飞行条件下表面所形成的激波形状和激波角,通过激波前后气流参数关系式计算波后气流参数,并利用二次激波膨胀波法结合修正牛顿理论确定飞行器表面边界层外缘气流参数,然后从二维平板低速层流的布拉修斯解出发,经相似比拟、Eckert参考焓法及高温气体特性等修正,推导得出的零攻角超声速球锥外形飞行器表面层流边界层内气动传热热流密度公式；另一方面建立并生成球锥外形飞行器模型及网格,将其导入程序中作为研究对象,并将推导得到的热流密度公式代入程序中作为计算模型的边界条件,然后采用有限体积法对控制方程(组)进行离散,最后通过迭代计算得到超声速飞行器气动热环境。本文主要研究内容有三部分：模拟球锥外形飞行器头部平面模型和三维模型在超声速飞行条件下的气动传热状况,结果表明在相同超声速飞行条件下,与平面模型相比三维模型产生的气动热环境更为严峻,且随着计算马赫数的增大,这种恶劣的热环境变得越来越严重；其次选取了三种不同半锥角的球锥外形飞行器模型,在相同计算条件下模拟结果显示,较高马赫数飞行条件下,对于头部曲率半径相同的模型,在合理范围内增大其半锥角,可以降低飞行器头部高温区域附近的温度值,即改善球锥外形飞行器在超声速飞行下恶劣的气动热环境；最后将飞行器模型内部空腔内表面和内空间环境之间的热传递考虑在内进行模拟计算,结果显示在较高的马赫数飞行条件下,考虑球锥外形飞行器内部封闭空腔内自然对流传热时,其所承受恶劣气动热环境得到了显著的改善。
[Abstract]:Based on the boundary layer theory, the classical empirical formula and numerical heat transfer coupling method are used to simulate the aerodynamic heat of a zero angle of attack supersonic spherical conical vehicle. The surface heat flux and the temperature distribution of the aircraft are obtained, and the aerodynamic thermal environment of the aircraft is analyzed accurately, which provides an important theoretical reference for the optimization of the shape design of the supersonic vehicle and the aerodynamic thermal design of the head. On the one hand, the flow field around the supersonic cone is solved to determine the shock wave shape and shock angle formed on the surface of the spherical cone shape aircraft under the supersonic flight condition, and the airflow parameters after the wave are calculated by the relation between the air flow parameters before and after the shock wave. Using the second shock expansion wave method and modified Newton theory, the outer boundary flow parameters of the aircraft surface boundary layer are determined. Then, based on the Brownian solution of the two-dimensional plate low speed laminar flow, the Eckert reference enthalpy method and the high temperature gas characteristics are compared with each other. The formula of aerodynamic heat transfer heat flux in laminar boundary layer on the surface of spherical conical shape vehicle with zero angle of attack is derived. On the other hand, the model and mesh of spherical cone shape aircraft are established and generated, and the model is introduced into the program as the object of study. The derived heat flux formula is added to the program as the boundary condition of the model, then the control equations are discretized by the finite volume method, and the aerodynamic thermal environment of the supersonic vehicle is obtained by iterative calculation. In this paper, there are three main parts: simulating the aerodynamic heat transfer of the spherical conical shape aircraft head plane model and three-dimensional model under supersonic flight conditions, the results show that under the same supersonic flight conditions, Compared with the plane model, the aerodynamic thermal environment generated by the 3D model is more severe, and with the increase of the Mach number, the bad thermal environment becomes more and more serious. Secondly, three kinds of spherical cone shape aircraft models with different semi-conical angles are selected. The simulation results under the same calculation conditions show that for the model with the same curvature radius of the head, the temperature near the high temperature region of the head can be reduced by increasing the semi-conical angle within a reasonable range under the condition of higher Mach number flight. That is to improve the bad aerodynamic and thermal environment of the spherical conical shape aircraft under supersonic flight. Finally, the heat transfer between the inner cavity surface and the inner space environment of the aircraft model is taken into account in the simulation calculation. The results show that when the natural convection heat transfer in the closed cavity of the spherical conical aircraft is considered under the condition of higher Mach number flight, the adverse aerodynamic thermal environment is greatly improved.
【学位授予单位】：南京理工大学
【学位级别】：硕士
【学位授予年份】：2015
【分类号】：V211

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