超音速绕凹角流动激波与紊流附面层干扰数值解

本文对超音速绕凹角激波与紊流附面层干扰流动进行了计算。计算采用Ce-beei-Keller Box方法;紊流模型用代数涡粘性模型;压强分布用流过尖劈统一的高超音速与超音速公式;对激波与紊流附面层干扰进行迭代修正。计算较好地预估了壁面压强分布以及压强开始升高点位置。     

跨音速粘流的计算

 概述了一种采用粘流/无粘流相互作用原理计算跨音速粘流的方法。采用熵修正激波算子计及跨越激波时的熵增,形成的非等熵位势方法可比传统的位势方法更准确地计算无粘流动。提出了一种流向速度型,结合其它辅助关系式导出了三维湍流边界层积分方程反方法。用此方法可求得粘流解。利用半反耦合方式耦合了无粘流和粘流解。数值算例表明,计算结果与实验结果吻合;且对计算机的要求较低。     

Hysteresis processes in the regular reflection↔Mach reflection transition in steady flows

Ernst Mach recorded experimentally, in the late 1870s, two different shock-wave reflection configurations and laid the foundations for one of the most exciting and active research field in an area that is generally known as Shock Wave Reflection Phenomena. The first wave reflection, a two-shock wave configuration, is known nowadays as regular reflection, RR, and the second wave reflection, a three-shock wave configuration, was named after Ernst Mach and is called nowadays Mach reflection, MR.A monograph entitled Shock Wave Reflection Phenomena, which was published by Ben-Dor in 1990, summarized the state-of-the-art of the reflection phenomena of shock waves in steady, pseudo-steady and unsteady flows.Intensive analytical, experimental and numerical investigations in the last decade, which were led mainly by Ben-Dor's research group and his collaboration with Chpoun's, Zeitoun's and Ivanov's research groups, shattered the state-of-the-knowledge, as it was presented in Ben-Dor (Shock Wave Reflection Phenomena, Springer, New York, 1991), for the case of steady flows. Skews's and Hornung's research groups joined in later and also contributed to the establishment of the new state-of-the-knowledge of the reflection of shock waves in steady flows.The new state-of-the-knowledge will be presented in this review. Specifically, the hysteresis phenomenon in the RR↔MR transition process, which until the early 1990s was believed not to exist, will be presented and described in detail, in a variety of experimental set-ups and geometries.Analytical, experimental and numerical investigations of the various hysteresis processes will be presented.     

跨声速翼型绕流计算边界层反方法的研究

本文用有限差分方法,通过有粘/无粘迭代计算了二元翼型的跨声速绕流问题。在边界层粘性区域内考虑了层流、转捩及湍流流动,当边界层内出现分离时,使用边界层反方法,采用代数湍流模型。算例表明,对激波/边界层弱干扰和强干扰情况,该方法的结果与风洞实验结果吻合良好,对于求解边界层小分离流场是一种好的近似方法。     

考虑粘性相互作用的跨音速机翼分析

本文将一些已有的三维全位势方程计算方法和一种三维积分边界层方法作了弱相互作用的耦合计算。计算了三种机翼并与已有的试验数据作了比较。     

Mach reflection in steady supersonic flow considering wedge boundary-layer correction

Mach reflection in steady supersonic flow is an important phenomenon having received extensive studies, among which simplified theoretical models to predict the size of Mach stem and other flow structure are of particular interest. Past efforts for such models were based on inviscid assumption while in real cases the flow is viscous. Here in this paper we consider the influence of wedge boundary layer on the Mach stem height. This is done by including a simplified boundary layer model into a recently published inviscid model. In this viscous model, the wedge angle and the trailing edge height, which control the Mach stem height, are replaced by their equivalent ones accounting for the displacement effect of the wedge boundary layer, with the boundary layer assumed to be laminar or fully turbulent. This viscous model is shown to compare well with numerical results by computational fluid dynamics and gives a Mach stem height as function of the Reynolds number and Mach number. It is shown that due to the viscous effect, the Mach stem height is increased, through increasing the effective wedge angle.     

密切内锥乘波前体进气道一体化设计和性能分析

采用特征线方法设计了具有直线初始激波、内收缩段消除激波反射、出口参数均匀可控的基准内锥流场。基于密切内锥(Osculating Inward turning Cone,OIC)乘波体设计方法,发展了密切内锥乘波前体进气道(Os-culating Inward turning Cone Waverider Inlet,OICWI)一体化设计技术。基于基准内锥流场和前体进气道一体化设计技术,设计了密切内锥乘波前体进气道。采用数值方法对设计的密切内锥乘波前体进气道进行了计算分析,结果表明无粘流场结构和基准内锥流场吻合,无粘模拟结果和理论设计结果吻合。粘性数值模拟结果显示一体化进气道具有较高的流量捕获率及总压恢复特性,进气道出口流场分布均匀。     

固体火箭喷管两相粘性跨音速流场计算

本文进行了固体火箭喷管两相粘性湍流跨音速流场计算.粘性的气相控制方程用隐式近似因子分解法求解,粒子方程采用跟踪粒子轨迹的特征线法求解,粘性湍流选用代数模型,气相和凝相充分地偶合.CFL数可以取500左右,收敛速度快,使粘性两相跨音速喷管流场计算耗费的机时达到工程计算可以接受的程度.通过计算,获得了在粘性和粒子同时作用下的流场参数分布.这对固体火箭发动机流场需要同时考虑粒子和粘性作用的专题研究大有裨益.     

超音速钝体无粘绕流的计算

研究用三维特征线相容性关系的迎风差分格式,数值模拟超音速和高超音速钝体无粘绕流流场。用钝体反方法计算头激波及其波后的流场,用轴对称流特征线法确定初值面上的流动参数。算例的计算结果与实验数据比较吻合     

A Structured Mesh Euler and Interactive Boundary Layer Method for Wing/Body Configurations

To compute transonic flows over a complex 3D aircraft configuration, a viscous/inviscid interaction method is developed by coupling an integral boundary-layer solver with an Eluer solver in a ＂semi-inverse＂ manner. For the turbulent boundary-layer, an integral method using Green＇s lag equation is coupled with the outer inviscid flow. A blowing velocity approach is used to simulate the displacement effects of the boundary layer. To predict the aerodynamic drag, it is developed a numerical technique called far-field method that is based on the momentum theorem, in which the total drag is divided into three component drags, i.e. viscous, induced and wave-formed. Consequently, it can provide more physical insight into the drag sources than the often-used surface integral technique. The drag decomposition can be achieved with help of the second law of thermodynamics, which implies that entropy increases and total pressure decreases only across shock wave along a streamline of an inviscid non-isentropic flow. This method has been applied to the DLR-F4 wing/body configuration showing results in good agreement with the wind tunnel data.     

Microphysics of Quasi-parallel Shocks in Collisionless Plasmas

Shocks in collisionless plasmas require dissipation mechanisms which couple fields and particles at scales much less than the conventional collisional mean free path. For quasi-parallel geometries, where the upstream magnetic field makes a small angle to the shock normal direction, wave-particle coupling produces a broad transition zone with large amplitude, nonlinear magnetic pulsations playing an important role. At high Mach numbers, ion reflection and acceleration are dominant processes which control the structure and dissipation at the shock. Accelerated particles produce a precursor, or foreshock, characterized by low frequency magnetic waves which are convected by the plasma flow into the shock transition zone. The interplay between energetic particles, waves, ion reflection and acceleration leads to a complicated interdependent system. This review discusses the spacecraft observations which have motivated the current view of the high Mach number quasi-parallel shock, and the theories and simulation studies which have led to a better understanding of the microphysics on which the quasi-parallel shock depends.     

不起动流场对超声速/高超声速进气道自起动性能的影响

对7个典型速域的二维超声速/高超声速进气道加速自起动过程进行了准定常数值仿真,分析了真实临界不起动流场对进气道自起动性能的影响,研究发现:存在一种介于超声速和高超声速临界不起动模式之间的过渡临界不起动模式。当真实不起动流场处于超声速临界不起动模式时,自起动马赫数略大于无黏设计自起动马赫数;处于过渡临界不起动模式时自起动马赫数小于无黏设计自起动马赫数;而该研究中处于高超声速临界不起动模式的进气道,自起动马赫数明显大于无黏设计自起动马赫数。高超声速临界不起动模式下的喉道截面特征气流参数显著偏离无黏临界不起动流场,所以Kantrowitz理论以及基于该理论发展而来的系列方法不适用于预测高超声速进气道自起动性能。     

Design method with controllable velocity direction at throat for inward-turning inlets

In the design of a hypersonic inward-turning inlet by applying the traditional basic flowfield, a reflected shock-wave is formed in the isolator due to the continuous reflection of the cowlreflected shock wave in the basic flow-field, which interacts with the boundary layer to produce a considerable influence on the performance of the inlet. Here, a basic flow-field design method that can control the velocity direction at the throat section is developed, and numerical simulations are conducted to demonstrate the effectiveness of this method. The method presented in this paper can achieve the absorption of the reflected waves at the shoulder of the basic flow-field by adjusting the variation law of the center radius in the basic flow-field, and a smooth transition between the compression surface and the isolator can also be produced. The Mach number and total pressure recovery coefficient of the inlet designed according to this method are 3.00 and 0.657, respectively, at design point(the incoming flow Mach number Ma1= 6.0). The results show that with this method, the inlet can efficiently weaken both the reflection of the shock wave and the interaction between the boundary layer and the reflected shock waves, which improves the aerodynamic performance of the inlet.     

高超音速平板前缘干扰数值解

本文采用了文献〔1〕中提出的一个数值求解Navier-Stokes方程的空间二阶精度单步差分格式,数值计算了超音速和高超音速平板前缘干扰问题。通过简单的模型方程对格式进行了描述和分析。用这一方法计算了M_∞=20,Tw/T_0=0.06,Re_(∞L)=4×10~3和M_∞=3,Tw=T_0,Re_(∞L)=10~3的平板前缘干扰问题。计算结果与实验值比较符合的很好,通过这些计算结果可使我们更清楚地了解平板前缘干扰的流动特性。     

一种基于三维粘性流场的乘波体生成方法

为了便于乘波体的优化设计工作,本文实现了一种基于任意三维粘性流场,在激波面上截取乘波体前缘线来生成乘波体外型的设计方法,并与NASA Langley研究中心的CFL3D软件的算例结果进行了比较检验,误差满足工程计算要求.由于在流场计算中直接计入了粘性,而不像大多数的研究方法那样通过无粘计算再进行粘性修正,而且对于高超声速乘波体设计,可以在程序中考虑真实气体效应、化学反应等因素,因此本文的设计方法更接近工程实际,更具有推广价值.     

圆形出口内转式进气道流动特征

采用数值仿真的方法研究了内转式进气道的流动特征。研究表明:设计状态在近壁面唇罩激波诱发了二次流,进而发展形成流向涡,造成低能流堆积,隔离段出口流场分布不均,消弱了进气道的抗反压能力。有攻角条件下,口面激波偏离唇罩前缘,激波形态发生改变,激波波面中部展向具有准二维特性,压缩面两侧气流压缩变弱,激波层变薄,出现局部膨胀区;有攻角条件下的无黏流场,在进气道压缩段形成三维流向涡,该流向涡促进高能高速气流向壁面迁移,改善了黏性条件下隔离段出口流场的均匀度。     

预处理方法在化学非平衡流场数值模拟中的应用

将预处理方法发展应用于化学非平衡流场的数值模拟中。采用有限差分,LU-SGS隐式方法对二维可压化学非平衡控制方程进行了耦合求解,对化学源项和无粘项隐式处理,粘性项则采用显式。通过化学反应剪切层和激波点火的算例表明,所采用的数值方法能够有效求解化学非平衡流场,证明了本算法的可行性,为下一步的工程应用奠定了基础。     

突然起动圆柱的非定常绕流计算

本文用分区方法定量模拟高Re数情况下突然起动圆柱的初期对称流动。绕圆柱的二维不可压流场可分为无粘区、附着粘性区和分离区。无粘区有解析解，用非定常层流边界层方程解附着粘性区，用离散涡方法处理随时间变化的分离区。试用了混合的Rankine和Lamb涡模型及二次离散涡。给出了Re=300、9500两个算例，基本符合有关实验结果。     

A study on turbulence transportation and modification of Spalart–Allmaras model for shock-wave/turbulent boundary layer interaction flow

It is of great significance to improve the accuracy of turbulence models in shock-wave/ boundary layer interaction flow. The relationship between the pressure gradient, as well as the shear layer, and the development of turbulent kinetic energy in impinging shock-wave/turbulent boundary layer interaction flow at Mach 2.25 is analyzed based on the data of direct numerical simulation(DNS). It is found that the turbulent kinetic energy is amplified by strong shear in the separation zone and the adverse pressure gradient near the separation point. The pressure gradient was non-dimensionalised with local density, velocity, and viscosity. Spalart–Allmaras(S–A) model is modified by introducing the non-dimensional pressure gradient into the production term of the eddy viscosity transportation equation. Simulation results show that the production and dissipation of eddy viscosity are strongly enhanced by the modification of S–A model. Compared with DNS and experimental data, the wall pressure and the wall skin friction coefficient as well as the velocity profile of the modified S–A model are obviously improved. Thus it can be concluded that the modification of S–A model with the pressure gradient can improve the predictive accuracy for simulating the shock-wave/turbulent boundary layer interaction.     

Comparison of CFD and theoretical post-shock gradients in hypersonic flow

In the recent work of Prof. Hans Hornung, expressions for the gradients of flow properties immediately behind a curved shock wave were obtained for a reacting gas [1]. In this paper, I use the expressions derived by Hornung to compare with inviscid computational fluid dynamics simulations of a Mach 8 flow over a cylinder. A finite-rate vibrational relaxation model is used to simplify the comparisons with theory. The shape of the bow shock wave is extracted from the CFD results, fitted with a polynomial, and then used to compute the post-shock gradients of the main flow variables. It is found that in general the CFD results are in very good agreement with the theory for both perfect gas and vibrationally relaxing flows. There are some notable differences, mostly centered on the location of the change in sign of the post-shock density gradient; this quantity is found to be highly sensitive to the relaxation rate of the gas. The theoretical post-shock gradients provide a rigorous test of CFD and suggest possible experiments that would be a very sensitive test of the models of finite-rate vibrational and chemical processes.