三维高超声速喷流干扰流场的数值模拟

采用ＬＵ隐式方法求解了有限体积法离散的完全Ｎ－Ｓ方程,数值模拟了三维高超声速喷流干扰流场。计算中采用了Ｊａｍｅｓｏｎ的当地极值递减（ＬＥＤ）格式和Ｂａｌｄｗｉｎ－Ｌｏｍａｘ代数湍流模型,与实验结果相比较,数值方法对附体流动计算较准确,对分离区附近流场计算有一定误差。     

“Ⅳ型”激波干扰中流-热-固耦合问题一体化计算分析

针对高超声速进气道前缘"Ⅳ型"激波干扰产生的气动加热与结构传热多物理场耦合计算问题，发展了一种基于有限体积法的流-热-固一体化计算方法。该方法采用一体化控制方程组统一离散求解外部高速流场与内部结构温度场，规避了传统分区耦合算法在时间域内交替迭代的繁琐数据交换策略。另外，提出一种新的双温阻模型计算流-固交界面的物性参数以保证计算准确性，采用LU-SGS隐式时间迭代和自适应时间步长以提高计算效率。采用经典高超声速二维圆管流-热-固耦合算例对该一体化方法进行验证，计算结果与试验值和参考文献数据吻合较好，证明了该方法的可靠性和正确性。利用一体化方法对高超声速前缘"Ⅳ型"激波干扰流-热-固耦合问题进行定常/非定常计算与分析，给出了温度与热流的时变特性，计算结果表明，激波干扰作用产生的超声速"喷流"不断冲击壁面，使得壁面最大压力系数增大约9倍，壁面最大热流增大约4.7倍，给高速飞行器的热防护设计与选材带来严峻挑战。同时，也表明了一体化计算方法可以较好地用于长航时飞行条件下与复杂飞行环境下的高超声速热防护系统的热环境特性分析与综合性能评估。     

Experimental and numerical study of the Mars Sample Return Orbiter in cold laminar hypersonic flow

This action has been undertaken in the framework of a collaborative program between NASA and CNES to support CNES in the development of the Mars Sample Return Orbiter (MSRO). The study, devoted to ONERA in this first phase of the program, consists in testing a 1/50^th size model in the hypersonic low Reynolds number wind tunnel R5Ch. Visualizations and heat-flux measurements have been carried out for different incidences in order to check the protection furnished to the satellite by the heat shield. The experimental measurements are discussed and compared with numerical results obtained with the Navier–Stokes finite volume solver FLU3M from ONERA. In addition, the influence on the measured heat flux of a partial cover, aiming at improving the thermal protection offered by the heat shield to the orbital vehicle is presented.     

Numerical simulation of multi-dimensional inviscid compressible flows by using TV-HLL scheme

This paper deals with the numerical solution of inviscid compressible flows. The threedimensional Euler equations describing the mentioned problem are presented and solved numerically with the finite volume method. The evaluation of the numerical flux at the interfaces is performed by using the Toro Vazquez-Harten Lax Leer(TV-HLL) scheme. An essential feature of the proposed scheme is to associate two systems of differential equations, called the advection system and the pressure system. It can be implemented with a very simple manner in the standard finite volume Euler and Navier–Stokes codes as extremely simple task. The scheme is applied to some test problems covering a wide spectrum of Mach numbers, including hypersonic, low speed flow and three-dimensional aerodynamics applications.     

Detached-Eddy Simulation of the vortical flow field about the VFE-2 delta wing

The numerical simulation of the flow for the VFE-2 delta wing configuration with rounded leading edges is presented using the Cobalt Navier–Stokes solver. Cobalt uses a cell-centered unstructured hybrid mesh approach, and several numerical results are presented for the steady RANS equations as well as for the unsteady DES and DDES hybrid approaches. Within this paper the focus is related to the dual primary vortex flow topology, especially the sensitivity of the flow to angle of attack and Reynolds number effects. Reasonable results are obtained with both steady RANS and SA-DDES simulations. The results are compared and verified by experimental data, including surface pressure and pressure sensitive paint results, and recommendations for improving future simulations are made.     

Numerical study of unsteady starting characteristics of a hypersonic inlet

The impulse and self starting characteristics of a mixed-compression hypersonic inlet designed at Mach number of 6.5 are studied by applying the unsteady computational fluid dynamics (CFD) method. The full Navier-Stokes equations are solved with the assumption of viscous perfect gas model, and the shear-stress transport (SST) k-x two-equation Reynolds averaged Navier- Stokes (RANS) model is used for turbulence modeling. Results indicate that during impulse starting, the flow field is divided into three zones with different aerodynamic parameters by primary shock and upstream-facing shock. The separation bubble on the shoulder of ramp undergoes a generating, growing, swallowing and disappearing process in sequence. But a separation bubble at the entrance of inlet exists until the freestream velocity is accelerated to the starting Mach number during self starting. The mass flux distribution of flow field is non-uniform because of the interaction between shock and boundary layer, so that the mass flow rate at throat is unsteady during impulse starting. The duration of impulse starting process increases almost linearly with the decrease of freestream Mach number but rises abruptly when the freestream Mach number approaches the starting Mach number. The accelerating performance of booster almost has no influence on the self starting ability of hypersonic inlet.     

Improved aerodynamic design of turbomachinery bladings by numerical optimization

The aerodynamic optimization of a transonic compressor is reported in this paper. The Q3D Navier–Stokes solver COLIBRI is coupled to a gradient-based method (CONMIN) and to a genetic algorithm (GADO). The suction side of a 2D blade is optimized by using both optimization methods with a significant efficiency improvement. In 3D, the performance improvement is obtained by modifying the suction surface of a transonic compressor with a Bézier surface and by using the CANARI solver coupled to the gradient method (CONMIN).     

Numerical simulation of roll induced moment of cruciform tactical missiles

This paper presents a numerical evaluation of the induced roll moment of cruciform missiles by a turbulent Navier–Stokes solver. The numerical results have been computed with a high resolution implicit upwind scheme. Comparisons with experiments performed at ONERA show a good agreement as well as a decisive improvement over numerical simulations based on Euler equations. Results are presented for subsonic and supersonic onflow conditions and two different configurations: fuselage-fins and fuselage-wing-fins. Comparative detailed analysis results in the origin of the rolling moment as well as the contributions of the different lifting surfaces.     

高超声速电离绕流的数值模拟

本文从三维N－S方程出发，采用稳式L－U分解算法和7组分15个反应的化学模型，数值模拟高超声速电离空气绕流。首先采用对称TVD格式、AUSMPW＋格式和Van Leer的矢通量分裂格式计算了高超声速球头绕流，并对它们的计算结果做了对比分析。然后用前两种格式，对RAM－C飞行试验模型三个再入高度（81km、71km、61km）的流场进行了数值模拟，计算的流场电子数密度值和试验测量数据符合较好。     

高速粘性绕流的有限体积TVD格式

本文通过求解Ｎａｖｉｅｒ－Ｓｔｏｋｅｓ方程,解决二维超声速及高速声速流动问题。通过采用坐标交换技术,可以方便地将有限体积ＴＶＤ格式应用到各种形状的流动问题中,各种典型算例的计算表明该方法数值计算稳定,捕捉激波分辩率高,能很好地模拟超声速和高速超声流动中激波反射及激波与边界层的干扰等各种复杂流动现象。湍流计算中采用ＢａｌｄｗｉｎＬｏｍａｘ湍流。数值结果与实验结果及理论解的对比令人满意。     

基于N-S方程的跨声速翼型多目标多约束优化设计

本文将粘性流场分析和数值优化方法耦合起来，发展了一种跨声速翼型设计方法，用以提高翼型在一个或多个设计点、在多种约束条件下的气动性能。由粘性流场分析程序计算得到的升力、阻力等气动参数构成目标函数，数值优化程序对其进行最小化。粘性流场分析采用了雷数平均Ｎ－Ｓ方程，这比过去翼型设计中使用的全速势方程或Ｅｕｌｅｒ方程更能模拟流动的本质，因而设计结果的可靠性大大提高了。优化算法采用传统的拟牛顿法（Ｑｕａｓｉ     

二维高亚声速空腔流激振荡的数值模拟研究

采用基于P .L .Roe的近似Riemann解的修正Osher&Chakravarthy (MOC)三阶TVD有限差分格式 ,数值求解二维雷诺平均全Navier Stokes方程 ,并用Cebeci Smith代数湍流模型 (对腔内区域作修正 )来模拟湍流效应 ,时间方向的积分采用四阶Runge Kutta方法 ,通过对GAMM超音速前台阶绕流的计算 ,验证了格式及程序的有效性 ,对高亚声速来流下的空腔流动作了数值模拟研究 ,取得了较好的效果     

导热-辐射耦合传热的多尺度分析和数值模型

复合材料高温传热特性的准确预测对飞行器热防护结构的设计有重要意义,相关模型也是国家数值风洞工程中多相多介质计算模型的重要组成部分。针对周期性结构复合材料的高温传热问题,利用多尺度渐进分析方法,对包含导热方程和辐射传输方程的导热-辐射耦合传热模型开展了研究。建立了表征单元尺度模型及宏观平均导热-辐射耦合传热模型,获得了材料宏观等效导热系数与表征单元模型间的关系,发现宏观等效辐射吸收和散射等系数可通过表征单元内的体积平均求取。结合有限容积方法与格子Boltzmann方法,建立了复合材料导热-辐射耦合传热多尺度数值模型。采用二维常物性材料传热过程的模拟验证了多尺度模型的有效性,结果表明所建立的多尺度模型能够对温度场给出准确高效的计算结果。该方法有助于为复合材料传热特性的预测提供数值手段。     

飞机增升装置气动力特性计算方法研究

本文给出了飞机增升装置气动力特笥的数值计算方法和计算结果，为了计算增升装置的缝隙效应和剪刀口效应，求解了粘性三维Navier-Stokes方程。为了正确 模拟大舵偏和阻力板后的大面积分离，采用了反映涡流特笥的新类型的两方程湍流模型。     

超燃发动机激波增强混合的数值研究

采用数值方法求解三维可压ＮＳ方程,模拟了斜激波增强超声速氢／空气混合的过程。对不同强度 激波混合增强的效率进行了比较。计算表明利用斜激波增强混合是一种行之有效的方法。     

有二次涡的激波边界层干扰流动的DPNS和NS方程计算

本文针对具有二次涡复杂分离再附现象的激波边界层干扰流动,数值地考察了扩散抛物化Ｎａｖｉｅｒ－Ｓｔｏｋｅｓ（ＤＰＮＳ）方程组的适用情况。壁面摩阻和压力、主涡和二次涡的涡高和涡长、分离再附位置以及流线图等特性的计算表明：ＤＰＮＳ方程组的数值结果均与ＮＳ方程组的数值结果很好相符。     

旋涡破裂引起的流动非定常特性

采用隐式近似因子分解法,通过求解三维、非定常、可压缩完全层流N-S方程,对绕大后掠三角翼旋涡破裂的流场进行了数值模拟。结果表明：当三角翼背风区流场发生旋涡破裂时,各流动参变量是非定常的,同时旋涡破裂点的位置沿轴向在一定范围内前后摆动。     

Numerical analysis on noise of rotor with unconventional blade tips based on CFD/Kirchhoff method

A solver is developed aiming at efficiently predicting rotor noise in hover and forward flight. In this solver, the nonlinear near-field solutions are calculated by a hybrid approach which includes the Navier-Stokes and Euler equations based on a moving-embedded grid system and adaptive grid methodology. A combination of the third-order upwind scheme and flux-difference splitting scheme, instead of the second-order center-difference scheme which may cause larger wake dissipation, has been employed in the present computational fluid dynamics (CFD) method. The sound pressure data in the near field can be calculated directly by solving the Navier-Stokes equations, and the sound propagation can be predicted by the Kirchhoff method. A harmonic expansion approach is presented for rotor far-field noise prediction, which gives an analytical expression for the integral function in the Kirchhoff formula. As a result, the interpolation process is simplified and the efficiency and accuracy of the interpolation are improved. Then, the high-speed impulsive (HIS) noise of a helicopter rotor at different tip Mach numbers and on different observers is calculated and analyzed in hover and forward flight, which shows a highly directional characteristic of the rotor HIS noise with a maximum value in the rotor plane, and the HSI noise weakens rapidly with the increasing of the directivity angle. In order to investigate the effects of the rotor blade-tip shape on its aeroacoustic characteristics, four kinds of blade tips are designed and their noise characteristics have been simulated. At last, a new unconventional CLOR-II blade tip has been designed, and the noise characteristics of the presented CLOR-II model rotor have been simulated and measured compared to the reference rotors with a rectangular or swept-back platform blade tip. The results demonstrate that the unconventional CLOR-II blade tip can significantly reduce the HSI noise of a rotor.     

双燃式冲压发动机中超燃燃烧室冷态流场数值模拟

采用Ｍａｃ Ｃｏｒｍ ａｃｋ 矢通分裂格式, 求解雷诺平均的Ｎ－Ｓ方程组, 采用Ｂａｌｄｗ ｉｎ－Ｌｏ－ｍ ａｘ 代数湍流模型和混合长度模型, 模拟超燃燃烧室内两股平行超声速来流的冷态流场, 取得了与试验相一致的结果, 计算表明回流区和射流流动参数对超燃燃烧室流场内波系的变化有着较大的影响。     

过冷水滴撞击结冰表面的数值模拟

为了计算过冷水滴撞击结冰表面,本文应用雷诺平均Navier-Stokes方程和k-ε两方程紊流模型计算结冰表面外的空气流场。用欧拉法建立了空气中过冷水滴的运动方程,并采用有限控制容积法对方程进行了数值计算。为了提高计算精度,方程中的对流—扩散项采用MUSCL格式进行了离散。计算结果表明过冷水滴的大小和机翼攻角的变化对结冰区的大小和水收集系数的影响很大。