Numerical Study on Hypersonic Flow and Aerodynamic Heating北大核心CSCD

基于有限差分法开发了高超声速流动与换热问题气热耦合仿真求解器,运用该求解器对三种典型高超声速流动与换热问题开展了仿真研究,得到了相应的气动参数、热流密度分布。高超声速后台阶的存在使表面气动参数、热流分布不再连续;随着缝深的提高,缝隙局部流速迅速降低,对流换热效应减弱;高超声速无限长圆管绕流中,边界层外部区域气动参数随时间变化不大,边界层内存在较大的温度梯度,壁面温度随时间升高。三个算例的仿真结果均与试验测量值进行了对比,验证了所开发的求解器的计算能力。     

零攻角小钝头钝锥高超音速绕流边界层的稳定性分析和转捩预报

研究了零攻角小钝头圆锥高超音速边界层的稳定性及转捩预测问题．小钝头的球头半径为0.5 mm,锥的半锥角为5°,来流马赫数为6．采用直接数值模拟方法得到了钝锥的基本流场,利用线性稳定性理论分析了等温壁面和绝热壁面条件下的第一、第二模态不稳定波,并用“e-N”方法对转捩位置进行了预测．在没有实验给出N值的情况下,暂取N为10．研究发现,壁面温度条件对于转捩位置有较大影响．绝热边界层的转捩位置比等温边界层的靠后．且尽管高马赫数下第二模态波的最大增长率远大于第一模态波的最大增长率,但绝热边界层的转捩位置是由第一模态不稳定波决定的．研究方法应能推广到有攻角的三维边界层流动的转捩预测．     

高超声速飞行器热防护材料与结构的研究进展

高超声速飞行器是航空航天的一个重要发展方向,在未来国防安全中起着重要作用．高超声速飞行器热防护材料与结构是高超声速飞行器设计与制造的关键技术之一,它关系到飞行器的安全．高超声速飞行器热防护材料与结构主要有金属TPS热防护系统、超高温陶瓷、C/C复合材料等．从材料制备、抗氧化、力学与物理性能表征等方面综述了热防护材料与结构的研究与应用现状,评述了其发展趋势．     

壁温变化非定常气动加热机理研究

研究了高超声速平板边界层因壁温时变引发的非定常气动热环境特征及机理.通过近似解析和数值模拟两种手段,得到了壁面热流随时问变化的完整过程.解析手段求解非定常可压缩边界层方程,将非定常响应表达成稳态解加上摄动级数的形式,在初始和最终稳态邻域分别求解,在适当的位置进行拼接,从而得到整个时间域上的解.在满足解析解假设的区域,数值结果与解析结果吻合较好,证明了所使用方法的可靠性.结果表明,非定常响应有两点显著特征:在壁温突然增加后短时间内,壁面热流方向改变,热边界层剖面在壁面附近出现了另一个拐点,这种新的剖面形状是典型的非定常特征.但是,高超声速情况下此种非定常响应存在的时间却很短,在考虑长时间气动加热的情况下,若只存在壁面温度时变的诱因,可以忽略流动中的非定常过程,当作准定常情况来处理.     

PSE在超音速边界层二次失稳问题中的应用

用抛物化稳定性方程（PSE）研究超音速边界层中的二次失稳问题．结果显示无论二维基本扰动是第一模态还是第二模态的T-S波,二次失稳机制都起作用．三维亚谐波的放大率随其展向波数和二维基本波幅值的变化关系与不可压缩边界层中所得类似．但是,即使二维波的幅值大到2％的量级,三维亚谐波的最大放大率仍远小于最不稳定的第二模态二维T-S波的放大率．因此,二次失稳应该不是导致超音速边界层转捩的主要因素．     

抛物化稳定性方程在可压缩边界层中应用的检验

用抛物化稳定性方程（PSE）,研究了可压缩边界层中扰动的演化,并与由直接数值模拟（DNS）所得进行比较．目的在检验PSE方法用于研究可压缩边界层中扰动演化的可靠性．结果显示,无论是亚音速还是超音速边界层,由PSE方法和由DNS方法所得结果都基本一致,而温度比速度吻合得更好．对超音速边界层,还计算了小扰动的中性曲线．与线性稳定性理论（LST）的结果相比,二者的关系和不可压边界层的情况相似．     

脉冲波作用下的高超音速非定常流动模拟及影响研究

采用高阶精度有限差分方法模拟了快声波脉冲扰动作用下的高超音速非定常流场,分析了脉冲波与高超音速流场的相互干扰,并应用Fourier频谱分析研究扰动波在边界层的发展．结果表明:来流脉冲扰动波与激波及边界层强烈相互作用,弓形激波明显向内弯曲,激波后扰动波被显著放大;来流扰动波与弓形激波干扰形成的边界层外的扰动波和近壁面内形成的边界层扰动波存在明显分界．钝锥头部参数扰动幅值要远大于其他位置参数扰动幅值．在边界层内的发展阶段,一些扰动模态持续增长,一些扰动模态被过滤掉,不再增长,甚至衰减,而也有一些扰动模态先衰减再增长．总的来说,在钝锥头部低频扰动模态为主导模态,随着扰动从流场上游向下游发展,总扰动模态中的低频模态成份和高频模态成份所占的比例开始转变,高频模态成分显著地增大．     

高超音速零攻角尖锥边界层转捩的机理

先计算出高超音速零攻角尖锥边界层的定常层流流场．然后在计算域的入口引入两组有限幅值的T-S波扰动,对空间模式的转捩过程进行了直接数值模拟．分析了转捩过程的机理,发现平均流剖面稳定性的变化是其关键．并进一步讨论了不同模态初始扰动在高超音速尖锥边界层中的演化规律．     

高超声速飞行器纵向动力系统平衡态集的拓扑结构研究

以高超声速飞行器纵向运动的空气动力学模型和结构动力学模型为依据,采用数值分析的方法,研究了高超声速飞行器动力系统平衡点集的拓扑结构.首先根据高超声速飞行器在巡航阶段的飞行边界的限制条件得到在给定的飞行高度和马赫数下的平衡点集,由此近似估算出了高超声速飞行器的飞行包线.然后根据得到的平衡点集,分别研究了高超声速飞行器的迎角、升降舵偏角和发动机的燃料当量比与飞行马赫数和高度的关系,并进行了数值拟合,在此基础上分别描绘了以上三个拟合关系式的曲面关系图。     

热传导对气-液射流界面不稳定性的作用机理研究

研究了平面分层气-液射流在非线性温度分布条件下的界面不稳定性性质.考虑了气体的可压缩性、液体的粘性、以及气体热导率和密度随温度变化等事实.并应用正则模态方法将问题转化为四阶变系数常微分方程,用数值积分和多重打靶法对模型的空间模式进行了计算,研究了不稳定模态随各物理参量的变化趋势.计算表明模型所体现的不稳定性特征与其它模型的计算结果是一致的.同时计算还得出气体和液体的温差越小、雷诺数越大、热导率变大均将有利于液体射流有效雾化的结果.该结论与HJE.Co.Inc(Glens Falls,NY,USA)的实验数据是定性吻合的.     

Thin-layer version of the moment equations in the boundary layer problem

The two-dimensional problem of a hypersonic kinetic boundary layer developing on a thin body in the case of a monatomic gas is considered. The model of the flow arises from the kinetic theory of gases and, within its accuracy, i.e., in the approximation of a hypersonic boundary layer, takes into account the strong nonequilibrium of the flow with respect to translational degrees of freedom. A method for representing the solution of the problem in terms of the solution of a similar classical (Navier-Stokes) hypersonic boundary layer problem is described. For the kinetic version of the problem, it is shown that the shear stress and the specific heat flux on the body surface are equal to their counterparts in the Navier-Stokes boundary layer.     

Mathematical model of a kinetic boundary layer

The two-dimensional (plane) problem of a hypersonic kinetic boundary layer developing on a thin body in the case of a homogeneous polyatomic gas flow with no dissociation or electron excitation is considered assuming that energy exchange between translational and internal molecular degrees of freedom is easy. (The approximation of a hypersonic kinetic boundary layer arises from the kinetic theory of gases and, within the thin-layer model, takes into account the strong nonequilibrium of the hypersonic flow with respect to translational and internal degrees of freedom of the gas particles.) A method is proposed for constructing the solution of the given kinetic problem in terms of a given solution of an equivalent well-studied classical Navier-Stokes hypersonic boundary layer problem (which is traditionally formulated on the basis of the Navier-Stokes equations).     

Nonequilibrium jet boundary layer in a polyatomic gas

The two-dimensional nonequilibrium hypersonic free jet boundary layer gas flow in the near wake of a body is studied using a closed system of macroscopic equations obtained (as a thin-layer version) from moment equations of kinetic origin for a polyatomic single-component gas with internal degrees of freedom. (This model is can be used to study flows with strong violations of equilibrium with respect to translational and internal degrees of freedom.) The solution of the problem under study (i.e., the kinetic model of a nonequilibrium homogeneous polyatomic gas flow in a free jet boundary layer) is shown to be related to the known solution of the well-studied simpler problem of a Navier-Stokes free jet boundary layer, and a method based on this relation is proposed for solving the former problem. It is established that the gas flow velocity distribution along the separating streamline in the kinetic problem of a free jet boundary layer coincides with the distribution obtained by solving the Navier-Stokes version of the problem. It is found that allowance for the nonequilibrium nature of the flow with respect to the internal and translational degrees of freedom of a single-component polyatomic gas in a hypersonic free jet boundary layer has no effect on the base pressure and the wake angle.     

The applicability of continuum models in the transitional regime of hypersonic flow over blunt bodies

Hypersonic rarefied gas flow over blunt bodies in the transitional flow regime (from continuum to free-molecule) is investigated. Asymptotically correct boundary conditions on the body surface are derived for the full and thin viscous shock layer models. The effect of taking into account the slip velocity and the temperature jump in the boundary condition along the surface on the extension of the limits of applicability of continuum models to high free-stream Knudsen numbers is investigated. Analytic relations are obtained, by an asymptotic method, for the heat transfer coefficient, the skin friction coefficient and the pressure as functions of the free-stream parameters and the geometry of the body in the flow field at low Reynolds number; the values of these coefficients approach their values in free-molecule flow (for unit accommodation coefficient) as the Reynolds number approaches zero. Numerical solutions of the thin viscous shock layer and full viscous shock layer equations, both with the no-slip boundary conditions and with boundary conditions taking into account the effects slip on the surface are obtained by the implicit finite-difference marching method of high accuracy of approximation. The asymptotic and numerical solutions are compared with the results of calculations by the Direct Simulation Monte Carlo method for flow over bodies of different shape and for the free-stream conditions corresponding to altitudes of 75–150 km of the trajectory of the Space Shuttle, and also with the known solutions for the free-molecule flow regine. The areas of applicability of the thin and full viscous shock layer models for calculating the pressure, skin friction and heat transfer on blunt bodies, in the hypersonic gas flow are estimated for various free-stream Knudsen numbers.     

Numerical study of the viscous hypersonic flow over a blunt delta wing

A previously developed two-dimensional Navier-Stokes solver is extended to three dimensions. The hypersonic steady flow of an perfect gas (with a constant adiabatic index) over a delta wing with a spherical nose and a cylindrical leading edge is computed as an example. The characteristic features of the flow are analyzed. Pressure, heat flux, and friction coefficient distributions are computed in the shock layer and on the wing surface. The results are compared with available numerical and experimental data.     

Effects of chemical reaction on MHD free convection and mass transfer flow of a micropolar fluid with oscillatory plate velocity and constant heat source in a rotating frame of reference

This paper concerns with studying the steady and unsteady MHD micropolar flow and mass transfers flow with constant heat source in a rotating frame of reference in the presence chemical reaction of the first-order, taking an oscillatory plate velocity and a constant suction velocity at the plate. The plate velocity is assumed to oscillate in time with a constant frequency; it is thus assumed that the solutions of the boundary layer are the same oscillatory type. The governing dimensionless equations are solved analytically after using small perturbation approximation. The effects of the various flow parameters and thermophysical properties on the velocity and temperature fields across the boundary layer are investigated. Numerical results of velocity profiles of micropolar fluids are compared with the corresponding flow problems for a Newtonian fluid. The results show that there exists completely oscillating behavior in the velocity distribution.     

Stability analysis of supersonic entropy layers including shock effects

The prediction of the laminar/turbulent transition location in supersonic boundary layers plays an important role to accurately compute aerodynamic forces and heating rates for the aerodynamic design and control of hypersonic vehicles. The stability characteristics of supersonic boundary layers depend e.g. on nose bluntness, transverse curvature, wall temperature, shock waves, etc. Most parameters can be theoretically investigated by performing conventional stability calculations with vanishing or asymptotic perturbation conditions at the far field. In this approach the formation of a shock in front of the leading edge of a blunt body is ignored. However, to improve the understanding of the interaction between instability waves originating inside supersonic boundary layer with those coming from the inviscid entropy layer, the presence of the shock has to be taken into account. This paper presents a method, how shock effects can be physically consistently included in stability calculations. The outer free‐stream boundary conditions are replaced by appropriate shock conditions. The required perturbation equations can be derived from the linearized unsteady Rankine‐Hugoniot equations, accounting for the effect of shock oscillations due to perturbated waves which originate from the flow field windward of the shock.     

Laws of the wall for velocity and temperature in supersonic turbulent boundary layers

Scaling laws for velocity and temperature profiles in the near-wall region of sub- and supersonic turbulent boundary layers have been developed, which allow us to represent velocity and temperature profiles in compressible gas stream in terms of those in an incompressible boundary layer. They are obtained as asymptotic expansions of the solutions to the Reynolds equations in a small parameter — the Mach number based on the friction velocity and gas enthalpy on the wall. The leading term of the expansion for velocity corresponds to known Van Driest's formula. However, the obtained solution contains additional terms of order unity, which explains the contradiction between Van Driest's formula and experimental data. The law of the wall for temperature, which has been formulated for the first time, has an analogous structure. Besides the von Kármán constant and the turbulent Prandtl number in the logarithmic region, known for incompressible flow, the obtained relations contain three new universal constants, which do not depend on gas molecular properties and the specific heat ratio. (© 2017 Wiley-VCH Verlag GmbH & Co. KGaA, Weinheim)