尖锐前缘气动加热机理与激波脱体距离研究

针对尖锐前缘所处的气动环境,分析了由于前缘尖化所带来的稀薄效应影响。应用理论方法推导了尖锐前缘驻点气动热的计算方法和计算公式,并结合CFD方法、DSMC方法和Fay-Riddel工程方法对其前缘钝头的流场、驻点热流进行了计算和对比分析。计算结果表明,由于前缘的尖化使得稀薄气体效应提前出现,从而影响到前缘的激波厚度、激波形状和激波脱体距离等流动现象,导致激波结构复杂化,可能会对进气道唇口的斜激波带来不利的影响。另外稀薄效应的影响降低了驻点热流,缓解了尖化前缘的气动热环境。     

基于多层分块算法的激波干扰流场预测

激波-激波干扰流场预测是超声速乃至高超声速流动中最具挑战性的问题之一. 特别地, 第IV类激波干扰由于其在壁面驻点附近产生极高的热载荷而备受关注. 本文针对圆柱诱导的弓形激波和入射斜激波的干扰问题, 分别基于量热完全气体模型和考虑振动激发的热完全气体模型, 数值求解有黏二维可压缩NS方程, 分析了高温气体效应对激波干扰流场结构, 以及第IV类激波干扰流场状态参数的影响. 接着, 本文基于一种具有广义可分离特性的遗传算法 (多层分块算法), 给出能够预测不同气体模型下第IV类激波干扰流场三波点的坐标位置、超声速射流的几何形状等特征性几何结构的数学模型, 进一步获得高温气体效应对激波干扰类型转变准则影响的定量化评估. 激波干扰类型转变准则面上的多组临界工况的激波干扰流场结构以及壁面压力和壁面热流分布的对比结果表明, 不同气体模型下的激波干扰类型和流场结构差异较为显著, 获得的定量化预测模型对工程中气动热环境的预测具有一定的参考价值.      

局部凸起在V形钝前缘模型中的降热特性研究

三维内转式进气道的唇口结构通常存在复杂的激波干扰及严酷的气动热载荷,严重威胁高超声速飞行器的性能与安全.在6.0马赫的高超声速流动中,以V形钝前缘模型为研究对象,设计了局部凸起的被动流动控制降热方案.采用数值模拟手段,首先研究了局部凸起方案的降热能力以及降热原理,然后初步优化了局部凸起的位置、高度以及宽度等关键设计参数,最后分析了优化后的局部凸起方案的攻角、侧滑角及马赫数的适用性.研究结果表明:上游凸起边缘形成的斜激波与主马赫反射结构形成的透射激波发生干扰,能够减弱其冲击壁面的强度,实现降热的目的;驻点凸起通过改变超声速射流的对撞角度,能够降低其对撞的强度,实现降热的目的.原始方案的降热能力约为37.75%,在对局部凸起的关键设计参数进行初步优化后,优化方案的降热能力将提升至44.60%.设计工况下的优化方案具有良好的攻角适用性,而高度可变的优化方案可以较好地适用于有侧滑角及高马赫数的流动.在研究范围内,高度可变的优化局部凸起方案的降热能力均高于20%.     

基于非线性耦合本构关系模型的尖化前缘气动加热影响研究

结合数值模拟与风洞试验技术,在高超声速连续/稀薄滑移流条件下对尖化前缘这一典型构型的气动加热影响开展深入研究.在三维有限体积框架下,应用非线性耦合本构关系(nonlinear coupled constitutive relations, NCCR)模型对试验工况下的尖化前缘外形开展数值计算,检验NCCR模型在尖化前缘构型中准确描述局部稀薄非平衡流动和物面气动热的性能.数值结果与实验数据对比表明,在等效高度33 km的风洞试验条件下, NCCR模型计算得到的驻点热流系数峰值同实验值偏差为1.81%, Fay-Riddell公式和纳维-斯托克斯(Navier-Stokes, NS)方程得到的驻点热流系数峰值同实验值偏差均在5%以内,物面其他位置的壁面热流系数计算值与实验值偏差均在10%以内,证明此时飞行器尖化前缘区域局部稀薄气体效应对气动加热影响程度较弱;在等效高度60 km时,飞行器尖化前缘区域附近的局部稀薄气体效应对气动加热的影响较为明显, NS方程计算的驻点热流系数偏差为33.31%, Fay-Riddell公式计算驻点热流系数同实验值偏差为29.5%, NCCR模型计算的驻点热流...     

极高超声速稀薄气体原子辐射效应的p-DSMC方法

极高超声速流动激波层内的高温导致内能模态的激发并伴随热辐射发生, 过高的温度使得空气分子完全解离, 原子组分对辐射热的贡献将达到80%以上. 本文基于优化的原子辐射模型, 提出追踪光子?直接模拟蒙特卡罗(p-DSMC)方法, 研究了稀薄流区不同马赫数下的高超声速二维圆柱绕流的壁面辐射加热, 获得了有无激发辐射效应的壁面压力和热流以及沿驻点线变化的平动、振动和转动温度. 在不考虑激发辐射效应的情况下, 得到的壁面压力和热流与已有的模拟结果符合的非常好, 误差均在5%以内, 尤其是在驻点位置, 误差在1%以内; 获得的平动、振动以及转动温度均与文献结果符合的很好. 在相同的来流条件下, 考虑辐射效应后发现, 来流速度低于10 km/s时, 辐射加热不明显, 在驻点区域, 辐射加热占对流加热比重在7%左右; 来流速度大于10 km/s时, 在驻点区域, 辐射加热占对流加热比重将超过30%. 考虑辐射效应后, 对非平衡区的平动、转动和振动温度的最大值影响不大. 此外, 另一个重要结论是, 流场中原子的浓度是影响壁面辐射热流大小的一个重要因素.      

环形激波聚焦流场特性的数值研究

针对环形激波聚焦过程产生的高温、高压特性,采用间断有限元方法模拟了环形激波在同轴圆柱 形激波管内的聚焦流场特性。计算结果表明,采用间断有限元方法能够有效地捕捉激波聚焦过程形成的二次 激波、涡环、三波交点和球面双马赫反射等主要流动特征。此外,通过改变环形管道内外半径对聚焦流场进行 模拟发现,环形管道外径对中心轴线上聚焦峰值压力的大小和位置影响较小,环形管道内径对中心轴线上聚 焦峰值压力的大小和位置影响较大。计算结果可以为工程应用提供一定的理论指导。     

高超声速钝前缘乘波构型优化设计研究

乘波体的高升阻比优势使其在高超声速飞行器设计中极具应用前景. 在实际工程应用中, 为了满足防热要求, 乘波体前缘必须进行钝化处理, 前缘钝化对乘波体气动性能会产生显著影响. 因此, 原始尖前缘最优乘波体并不能保证钝化后仍为最优. 针对这一问题, 首先研究了前缘钝化对不同构型升阻特性的影响程度和作用机理. 结果表明: 前缘钝化会造成乘波体升力小幅度降低, 阻力大幅增加, 升阻比显著降低; 其中钝前缘本身的波阻在阻力增量中起主导作用, 而钝前缘本身的摩阻增加量与物面的摩阻降低量十分接近. 基于上述结果, 提出了一种高效评估钝前缘乘波体气动力的方法, 并结合遗传算法, 开展了直接考虑前缘钝化影响的乘波构型优化设计研究, 获得了钝前缘最优构型. 通过CFD数值模拟对最优构型的气动力特性进行评估, 结果表明: 在不同飞行高度、不同升力和不同钝化半径约束下, 相比尖前缘最优构型, 钝前缘最优构型宽度变窄, 相同纵向位置处的后掠角增大, 且升阻比显著提升. 在M_∞ = 15, H = 50 km, CL = 0.3约束条件下, 钝化半径R = 10 mm的钝前缘最优构型设计点升阻比相比尖前缘最优构型提升量可达9.32%.      

进气道等离子体/磁流体流动控制研究进展

为实现高速飞行器的宽速域飞行,如何保证进气道在非设计状态下的性能至关重要。相比于传统被动控制方式,等离子体/磁流体流动控制技术作为新概念主动流动控制技术,由于其具有结构简单,快速响应,并可根据实际飞行条件进行反馈控制等优势,在国内外上得到了广泛关注。本文介绍了等离子体/磁流体在高超/超声速进气道的主要应用方式与等离子体/磁流体建模方法。当进气道处于超临界状态时,等离子体/磁流体流动控制主要通过热阻塞效应产生虚拟型面,从而将激波系推回至唇口,该技术有望在需要短时间流动控制的高马赫数导弹上走向工程应用;由于等离子体/磁流体激励器与壁面平齐安装,对于高超声速飞行条件,相比于粗糙元其对热防护的要求较低,并且通过超声速风洞实验初步证明了通过高频激励对边界层施加扰动的可行性,需要从稳定性理论的角度对其物理机制进行研究。在后续发展中需要进一步创新等离子体产生技术及激励方式,发展等离子体与流的全耦合计算模型等离子体与流的全耦合计算模型与高效算法 ,为指导工程应用提供依据.      

稀薄流非线性模型方程离散速度坐标法有限差分解

从一般非线性Bo ltzm ann方程出发,发展并实现了一套适于大范围K nudsen数稀薄流问题数值模拟的统一算法。采用BGK模型和Shakov模型近似碰撞项,进而引入两个二速度无量纲简化分布函数,通过关于分子速度第三分量取矩积分,将三速度单一模型方程变换为二速度微分方程组。基于G auss-H erm ite积分公式和正交多项式G auss积分公式,借助离散速度坐标法消除简化模型方程对分子速度空间的连续依赖性,从相空间到物理空间得到一组带源项双曲守恒离散方程,并给出其显式和隐式二阶迎风TVD有限差分解。以二维圆柱A r气体超声速绕流算例,验证了数值算法的有效性,比较分析了漫反射和镜面反射两种气体分子壁面反射模型的计算结果。     

垂直壁面附近上升单气泡的弹跳动力学研究

采用高速摄影技术结合阴影法,对静止水中垂直壁面附近上升单气泡运动进行实验研究,对比气泡尺度及气泡喷嘴与壁面之间的初始无量纲距离(S~*)对气泡上升运动特性的影响,分析气泡与壁面碰撞前后,壁面效应与气泡动力学机制及能量变化规律.结果表明,对于雷诺数Re≈580～1100,无量纲距离S~*2～3时,气泡与壁面碰撞且气泡轨迹由无约束条件下的三维螺旋转变成二维之字形周期运动;当S~* 2～3时,壁面效应减弱,有壁面约束的气泡运动与无约束气泡运动特性趋于一致.气泡与壁面碰撞前后,壁面效应导致横向速度峰值下降为原峰值的70%,垂直速度下降50%;气泡与壁面碰撞前,通过气泡中心与壁面距离(x/R)和修正的斯托克斯数相关式可预测垂直速度的变化规律.上升气泡与壁面碰撞过程中,气泡表面变形能量单向传输给气泡横向动能,使得可变形气泡能够保持相对恒定的弹跳运动.提出了气泡在与壁面反复弹跳时的平均阻力系数的预测模型,能够很好地描述实验数据反映出的对雷诺数Re、韦伯数We和奥特沃斯数Eo等各无量纲参数的标度规律.     

NUMERICAL SIMULATION OF THREE-DIMENSIONAL SHOCK-SHOCK INTERACTIONS ON A BLUNT BODY

A numerical study is conducted to simulate the effects of extraneous shock impingement on a blunt body in viscous hypersonic flow. The interaction of extraneous shock with the leading-edge shock results in a very complex flow field that contains local regions of high pressure and intense heating. The heating and pressure can be orders of magnitude higher than the peak values in the absence of shock impingement. The flow field is calculated by solving thin-layer Navier-Stokes equations with a finite-volume flux splitting technique developed by van Leer. For a zero or small sweep of the body, a type IV interaction occurs, which produces a lambda shock structure with a supersonic jet embedded in the otherwise subsonic flow; for a moderate sweep of about 25^°, a type V interaction occurs in which a subsonic shear layer sandwiched in supersonic flow is produced with a transmitted shock. In the present study, both type IV and type V interactions are investigated. Results of the present numerical investigation are compared with available experimental results. For the present conditions, the peak pressure is 2.2 times the unimpinged stagnation point pressure and the peak heating is 3 times the unimpinged stagnation point heating. The flow for a type IV interaction is found to be unsteady.     

Numerical heat transfer study around a spiked blunt-nose body at Mach 6

An aerospike attached to a blunt body significantly alters its flowfield and influences aerodynamic drag at high speeds. The dynamic pressure in the recirculation area is highly reduced and this leads to the decrease in the aerodynamic drag. Consequently, the geometry of the aerospike has to be simulated in order to obtain a large conical recirculation region in front of the blunt body to get beneficial drag reduction. Axisymmetric compressible Navier–Stokes equations are solved using a finite volume discretization in conjunction with a multistage Runge–Kutta time stepping scheme. The effect of the various types of aerospike configurations on the reduction of aerodynamic drag is evaluated numerically at a length to diameter ratio of 0.5, at Mach 6 and at a zero angle of incidence. The computed density contours are showing satisfactory agreement with the schlieren pictures. The calculated pressure distribution on the blunt body compares well with the measured pressure data on the blunt body. Flowfield features such as formation of shock waves, separation region and reattachment point are examined for the flat-disc spike and on the hemispherical disc spike attached to the blunt body. One of the critical heating areas is at the stagnation point of a blunt body, where the incoming hypersonic flow is brought to rest by a normal shock and adiabatic compression. Therefore, the problem of computing the heat transfer rate near the stagnation point needs a solution of the entire flowfield from the shock to the spike body. The shock distance ahead of the hemisphere and the flat-disc is compared with the analytical solution and a good agreement is found between them. The influence of the shock wave generated from the spike is used to analyze the pressure distribution, the coefficient of skin friction and the wall heat flux facing the spike surface to the flow direction.     

高超声速飞行器气动防热新概念研究

传统乘波构型的高超声速飞行器尖锐的前缘存在严重的气动加热问题,而简单的前缘钝化气动防热方法由于造成很大的升阻比损失,难以发挥实质性作用. 引入``人工钝前缘(ABLE)'概念,拟以一种新的思路解决这一矛盾. 通过定义ABLE构型的外形参数,并采用CFD数值计算方法研究了各参数对气动力和气动热特性的影响规律,在流场分析的基础上进行了外形优化,最终得到令人满意的新型高超声速飞行器头部外形,总结了运用ABLE概念进行气动防热的相关设计原则和规律.      

Transverse and parallel injection of hydrogen with supersonic combustion in a shock tunnel

An experimental comparison has been made of the combustion induced pressure rise in a constant area duct when hydrogen is injected transverse to the flow by using a surface orifice, and when it is injected parallel to the flow by using a central injection strut. The experiments were conducted in a shock tunnel at a flow Mach number of 4.2 and stagnation enthalpies of 5.6, 6.5 and 8.9 MJ kg. Both room temperature and heated hydrogen were injected, and a method of heating the hydrogen by compression in a gun tunnel which was slaved to the shock tunnel is described. It was found that, for both unheated and heated hydrogen, the combustion pressure rise was not measurably dependent on the method of introducing the hydrogen, not withstanding the complicated shock related flow pattern arising from transverse injection. Received August 14, 1995 / Accepted February 14, 1996     

圆球诱发斜爆轰波的数值研究

斜爆轰发动机是飞行器在高马赫数飞行条件下的一种新型发动机,具有结构简单、成本低和比冲高等优点.但是斜爆轰发动机的来流马赫数范围广,来流条件复杂,为实现斜爆轰波的迅速、可靠引发,采用钝头体来诱发.利用Euler方程和氢氧基元反应模型,对超声速氢气/空气混合气体中圆球诱导的斜爆轰流场进行了数值研究.不同于楔面诱发的斜爆轰波,球体首先会在驻点附近诱发正激波/爆轰波,然后在稀疏波作用下发展为斜激波/爆轰波.模拟结果显示,经过钝头体压缩的预混气体达到自燃温度后,会出现两种流场:当马赫数较低时,由于稀疏波的影响,燃烧熄灭,钝头体下游不会出现燃烧情况;而当马赫数较高时,燃烧阵面能传到下游.分析表明,当钝头体的尺度较小时,驻点附近的能量不足以诱发爆轰波,只会形成明显的燃烧带与激波非耦合结构;当钝头体的尺度较大时,流场中不会出现燃烧带与激波的非耦合现象,且这一特征与马赫数无关.通过调整球体直径,获得了激波和燃烧带部分耦合的燃烧流场结构,这一流场结构在楔面诱发的斜爆轰波中并不存在,说明稀疏波与爆轰波面的相互作用是决定圆球诱发斜爆轰波的关键.     

High enthalpy nozzle flow sensitivity study and effects on heat transfer

The sensitivity of the flow along the nozzle and in the test section of high enthalpy wind tunnels to the thermochemical response of the nozzle expansion process, as well as effects on the pressure and heat transfer distributions over the Electre blunt cone standard test model, are examined in the framework of properly characterizing the test section flow field in such facilities. Particularly sensitive to the thermochemical behaviour of the nozzle flow, in the facilities under consideration, are the static pressure, static temperature and Mach number, whereas stagnation point (pitot) pressure and heat transfer data or freestream velocity are inadequate for the characterization of the thermochemical state of the flow. The Electre and nozzle wall pressure data in the F4 arc jet wind tunnel suggest, in contrast to nonequilibrium computations, that the flow in the F4 nozzle is close to equilibrium. In the HEG and, to some extent, the T5 piston-driven shock tunnels, there are indications that significant heat losses occur in the reservoir. Lastly, simple semi-empirical formulations for stagnation point heating are shown to perform reasonably well in high enthalpy flow conditions.     

Numerical study of the oscillations induced by shock/shock interaction in hypersonic double-wedge flows

In this paper, the shock pattern oscillations induced by shock/shock interactions over double-wedge geometries in hypersonic flows were studied numerically by solving 2D inviscid Euler equations for a multi-species system. Laminar viscous effects were considered in some cases. Temperature-dependent thermodynamic properties were employed in the state and energy equations for consideration of the distinct change of the thermodynamic state. It was shown that the oscillation results in high-frequency fluctuations of heating and pressure loads over wedge surfaces. In a case with a relatively lower free-stream Mach number, the shock/shock interaction structure maintains a seven-shock configuration during the entire oscillation process. On the other hand, the oscillation is accompanied by a transition between a six-shock configuration (regular interaction) and a seven-shock configuration (Mach interaction) in a case with a higher free-stream Mach number. Numerical results also indicate that the critical wedge angle for the transition from a steady to an oscillation solution is higher compared to the corresponding value in earlier numerical research in which the perfect diatomic gas model was used.      

Shock tunnel study of spiked aerodynamic bodies flying at hypersonic Mach numbers

A three-component accelerometer balance system is used to study the drag reduction effect of an aerodisc on large angle blunt cones flying at hypersonic Mach numbers. Measurements in a hypersonic shock tunnel at a freestream Mach number of 5.75 indicate more than 50% reduction in the drag coefficient for a 120° apex angle blunt cone with a forward facing aerospike having a flat faced aerodisc at moderate angles of attack. Enhancement of drag has been observed for higher angles of attack due to the impingement of the flow separation shock on the windward side of the cone. The flowfields around the large angle blunt cone with aerospike assembly flying at hypersonic Mach numbers are also simulated numerically using a commercial CFD code. The pressure and density levels on the model surface, which is under the aerodynamic shadow of the flat disc tipped spike, are found very low and a drag reduction of 64.34% has been deduced numerically.     

AN APPROXIMATE MODEL OF UNSTEADY AERODYNAMICS FOR HYPERSONIC PROBLEMS AT HIGH ALTITUDE

针对高空高马赫数飞行环境和强黏性干扰的物理特性, 在当地流活塞理论的基础上引入有效外形修正, 发展了黏性修正当地流活塞理论, 结合定常N-S方程解给出了高空高马赫数下针对该方法的有效外形的判据, 并通过数值算例对该判据进行了验证.通过对典型尖头薄翼和典型钝头翼的一系列二维非定常算例, 将该方法与一阶活塞理论、基于欧拉(Euler)方程的当地流活塞理论和非定常N-S方程数值解进行了对比. 结果显示在高度为40~70 km、马赫数为10~20范围内, 通过该方法计算得到的非定常气动力与非定常N-S方程数值解吻合较好, 明显优于活塞理论和基于Euler方程的当地流活塞理论.该方法克服了传统的活塞理论和当地流活塞理论不能用于高空高马赫数这类强黏性效应情况的弊端, 在较宽的马赫数、攻角、飞行高度范围内都有良好的适用性, 同时其计算效率远高于非定常N-S方程.