近似黎曼解对高超声速气动热计算的影响研究

高超声速流场计算一般采用TVD型格式,这些格式中,大多采用了不同形式的近似黎曼解. 通过分析和数值验证,论述了激波捕捉格式中近似黎曼解的耗散性质,说明其对高超声速热流计算的影响. 数值实验证明,采用低耗散格式可大大提高热流计算精度,降低热流计算对网格的依赖程度,从而获得精确的热流数值解.      

旋成体表面防热的发汗冷却控制及其数值模拟

研究高超声速旋成体表面防热的发汗冷却控制系统,除去得到与平面情形相当的临界发汗通量之外,还得出另一个较小的第二临界发汗通量.当发汗通量介于这两个临界发汗通量之间时,表面虽将出现烧蚀,但烧蚀会自动停止,留有剩余厚度.还进行了数值模拟,给出各种特性曲线;详细讨论了对热层表面烧蚀控制的三种方式:表面温度的控制、表面烧蚀量的控制和表面烧蚀开始时间的控制,给出控制变量选取的准则.     

Developing a hybrid flux function suitable for hypersonic flow simulation with high‐order methods

In this paper, we develop a new hybrid Euler flux function based on Roe's flux difference scheme, which is free from shock instability and still preserves the accuracy and efficiency of Roe's flux scheme. For computational cost, only 5% extra CPU time is required compared with Roe's FDS. In hypersonic flow simulation with high‐order methods, the hybrid flux function would automatically switch to the Rusanov flux function near shock waves to improve the robustness, and in smooth regions, Roe's FDS would be recovered so that the advantages of high‐order methods can be maintained. Multidimensional dissipation is introduced to eliminate the adverse effects caused by flux function switching and further enhance the robustness of shock‐capturing, especially when the shock waves are not aligned with grids. A series of tests shows that this new hybrid flux function with a high‐order weighted compact nonlinear scheme is not only robust for shock‐capturing but also accurate for hypersonic heat transfer prediction. Copyright © 2015 John Wiley & Sons, Ltd.     

高导C/C材料在疏导式热防护中的应用探索

针对高超声速飞行器飞行时气动加热严重的问题，为了保证高升阻比外形，提出疏导式热防护结构，建立了一套内置高导C/C材料的疏导式热防护结构原理模型，通过数值模拟和电弧风洞的方法对疏导式热防护结构进行了分析，得到内置高导C/C材料的防热效果．数值模拟结果表明来流马赫数为8时，模型驻点温度下降了500度，柱面最低升高了380度，实现了热流从高温区到低温区的疏导，减弱了端头的热载荷，强化了端头的热防护能力．通过电弧风洞试验可以获得相似的结果，内置普通C/C材料表层抗氧化层出现严重烧蚀，而内置高导C/C材料基本不变，验证了数值模拟方法的准确性以及内置高导C/C材料疏导式热防护结构的有效性．     

An enthalpy‐preserving shock‐capturing term for residual distribution schemes

In this contribution, we investigate strategies to perform shock‐capturing computation of steady hypersonic flow fields by means of residual distribution schemes. The ultimate objective is the computation of flow solutions for which the correct upstream enthalpy value is recovered in the postshock region. To this end, the parallelism existing between the classical Bx scheme and the stabilized finite element techniques is exploited. The simple Lax‐Friedrichs dissipation term is leveraged to build two new residual distribution schemes. Upon testing on both inviscid and viscous steady problems, solutions obtained with one of the two schemes are shown to recover the correct upstream total enthalpy level in the postshock region. This last scheme provides also improved wall pressure and skin friction predictions; heat transfer predictions are, unfortunately, similar to those offered by the Bx scheme. A conjecture for explaining this behavior is exposed.     

THE INVESTIGATION OF EXPERIMENTAL TECHNIQUE FOR HIGH TEMPERATURE GAS JET FLOW TEST IN IMPULSE WIND TUNNEL

报道关于高温燃气自由喷流（热喷流）、燃气喷流／主流干扰流对气动热环境影响的实验研究结果．其意义在于：抽象出高超声速飞行器实际飞行时燃气喷流及其干扰流的物理模型，为高超声速飞行器防热需求提供实验依据．实验主流由脉冲风洞提供，燃气喷流用氢氧燃烧驱动路德维希管的方式产生．利用脉冲风洞驱动段压力信号自动控制热气源的产生以保证风洞主流与燃气喷流同步，利用氢气、氮气和氧气的不同比例实现燃气喷流的热力学相似．实验技术上完成了高温燃气喷流系统的参数采集与系统状态标定；实验内容上开展了压缩拐角平板模型的气动热实验研究，通过实验比较了只有主流流场、只有热喷流流场和既有主流流场又有热喷流流场（即干扰流场）3种工况的热流分布．实验研究发现，热喷流／主流相互干扰会对压缩拐角平板上某一范围内的气动热环境造成显著影响，热流峰值较无喷流流场高出一个量级．     

再入飞行器湍流尾迹流场研究

再入飞行器湍流尾迹流场状况，直接关系到飞行器的雷达散射特性。对再入飞行器湍流尾迹等离子体场理论模型，试图通过湍流模式理论来表达，即使用κ－ε－g模型方程来封闭平均化的全Navier-Stokes方程，从而准确获得流动平均场和脉动场信息。使用的N－S平均方程由质量加权平均过程产生，湍流模型方也经过可压缩性修正。真实气体效应重点考察空气处于局部热化学平衡状态。流动控制方程运用一个二阶TVD格式的有限体积法求解，以一典型小钝锥体零攻角再入飞行为例，计算了在两种高程（H－40km和H＝30km)条件下的高超声速湍流尾迹流场。获得的尾迹流场参数与流动物理状况符合，并且湍流脉动参数与已有相应的实验结果定性一致，初步证实该方法合理。     

Evolution of Controlled Disturbances in the Shock Layer on the Compression Surface

Results of an experimental study of density-wave characteristics in the shock layer are presented for the case of a hypersonic nitrogen flow around a model with a two-dimensional compression surface, which is an arc of a circle, and a sharp leading edge. Controlled periodic disturbances developed on the streaky structure are registered by the electron-beam fluorescence technique. The streaky structure of the type of two vortices rotating in the opposite directions is generated in the shock layer by an oblique gas-dynamic whistle.     

Development of slow disturbances in a hypersonic boundary layer

The mechanisms of development of slow time-dependent disturbances in the wall region of a hypersonic boundary layer are established and a diagram of the disturbed flow patterns is plotted; the corresponding nonlinear boundary value problem is formulated for each of these regimes. It is shown that the main factors that form the disturbed flow are the gas enthalpy near the body surface, the local viscous-inviscid interaction level, and the type, either subsonic or supersonic, of the boundary layer as a whole. Numerical and analytical solutions are obtained in the linear approximation. It is established that enhancement of the local viscous-inviscid interaction or an increased role for the main supersonic region of the boundary layer makes the disturbed flow by and large “supersonic”: the upstream propagation of the disturbances becomes weaker, while their downstream growth is amplified. Contrariwise, local viscous-inviscid interaction attenuation or an increased role for the main subsonic region of the boundary layer has the opposite effect. Surface cooling favors an increased effect of the main region of the boundary layer while heating favors an increased wall region effect. It is also found that in the regimes considered disturbances travel from the turbulent flow region downstream of the disturbed region under consideration counter to the oncoming flow, which may be of considerable significance in constructing the nonlinear stability theory.     

Asymptotic solution of the two-dimensional equations for a thin viscous rarefied shock layer on a cold surface

Hypersonic rarefied flow past blunt bodies is studied in the continuum-free-molecular transition regime. On the basis of an asymptotic analysis three rarefied gas flow patterns are established depending on the relation between the relevant parameters of the problem. In the first regime corresponding to a cold surface asymptotic solutions of the equations of a thin viscous shock layer are derived at low Reynolds numbers in the axisymmetric and plane cases. Simple analytical expressions for the pressure and the heat transfer and friction coefficients are obtained as functions of the freestream parameters and the body geometry. With decrease in the Reynolds number the coefficients approach the values corresponding to free-molecular flow. In this regime a similarity parameter for the hypersonic rarefied flow past bodies is determined. The asymptotic solutions are compared with numerical solutions and the results of direct statistical simulation by the Monte Carlo method.