利用非结构化网格方法对翼型绕流的数值研究

利用同位非结构化网格上的压力加权修正算法 ,对翼型湍流绕流进行了数值分析。详细地给出了一孤立翼型在不同攻角下的分离流结构及翼型表面压力分布 ,为了显示非结构化网格方法在求解多连通流动区域的优越性 ,对双翼型绕流进行了数值计算。在数值分析中 ,对阵面推进法进行改进来生成三角形网格 ,采用有限控制体方法直接在物理空间中的非结构化网格单元上离散 Navier- Stokes方程及 k- ε方程 ,形成的代数方程组通过预条件矩阵共轭梯度平方法求解。计算结果表明 :当流动为附着流时 ,计算结果与实验值吻合程度令人相当满意 ;而在分离区内 ,计算结果与实验值存在一定的误差 ,需对分离区内的湍流模型做进一步的改进。     

绕流直接数值模拟误差分析

对于无边界绕流问题的计算流体力学模拟通常是将物体置于“足够大”的槽道中,而通过不断改变槽道尺寸以及离散网格密度,后验对比方式来检查模拟误差。本文结合多种经典流场理论,提出一种简单的先验误差估计方法来确定槽道尺寸以及相应的网格分布。在此方法中,对于槽道尺寸的确定基于线性叠加原理（即在极小雷诺数下采用Stokes理论解叠加,而在其他雷诺数条件下采用势流理论解叠加）,来估计槽道尺寸对绕流结果的影响。而对网格尺寸与分布,则是使用多项式逼近中的基本误差分析工具,应用到速度边界层,远场势流,以及Rankine涡等简单流动,从而确定整个绕流问题中的离散误差。为了验证前面的理论分析结果,本文模拟了相当大雷诺数范围内的二维翼型以及三维圆球绕流,所得数值结果非常好地验证了理论分析。结果表明,对于Stokes流动问题,槽道尺寸需要大约100倍于物体特征尺寸来保证其结果与无边界绕流相差不超过1%;而在雷诺数超过大约100时,槽道尺寸只需10倍（二维绕流）或者5倍（三维绕流）于物体特征尺寸来达到同等精度。在此先验误差估计方法可应用于一般化的绕流问题。

     

基于分区拼接网格技术高升力装置流场数值模拟

针对高升力装置构型模型结构复杂、流场变化剧烈等特点,本文采用分区拼接网格的思想分别按照流场和结构拓扑特点对高升力装置进行了网格分区。在分区的基础上逐块生成网格,减小了增升装置网格的生成难度,提高了网格质量,减少了网格数目。首先,研究了高升力装置的分区策略及流场特点;接着以MD30P-30N多段翼型为研究算例研究了网格比例和插值方法对计算结果的影响。经过分析对比可知:外部区域与近壁区域之间的比例不宜小于1:5;内部域网格比例不宜超过1:1.8,最好保持在1:1左右;计算中应该采用高阶精度插值以保证计算精度。采用某四段翼型进行了验证;最后采用NASA标准高升力装置进行了三维高升力装置流场数值模拟并与相应风洞实验数据及对接网格计算结果进行了比较与分析,验证了拼接网格技术的高效性与可靠性。同时分析研究了绕三维增升装置的流动及其周围复杂的粘性流动现象。     

多段翼型缝翼流动速度的定常和非定常特性研究

用热线风速仪研究多段翼型前缘缝翼在不同条件下流动速度的定常性和非定常性。结合多段翼型定常流动Navier-Stokes方程数值模拟的结果,分析了迎角、后缘襟翼参数(偏角、缝道宽度、搭接量)对缝翼定常和非定常流动速度的影响规律。研究结果表明:在缝翼后缘处,流动分为缝道加速流动区、缝翼尾流区、缝翼上表面以上的主流区;缝翼尾流区流动速度非定常性主要表现在中低频率范围(2k Hz以下),而缝道加速流动区和缝翼上表面以上的主流区流动非定常性常表现出高频特性(2k Hz以上);在失速前随迎角增加,或者当襟翼偏角从20°向30°增加时,缝道流动加速,槽区涡减小;缝翼槽区涡形成和振荡是中低频率范围流动非定常性的机理,而缝翼鼻尖脱落涡是缝翼槽区涡振荡的激励因素。     

THE NUMERICAL STUDY OF LATTICE BOLTZMANN METHOD BASED ON DIFFERENT GRID STRUCTURE

传统的格子波尔兹曼方法(lattice-Boltzmann method, LBM)通常基于标准均匀网格, 这主要取决于速度的空 间离散格式.均匀网格结构的特点, 使LBM在处理具有复杂边界的问题时遇到较大的困难, 从而限制了它的应用.另外, 对于较为复杂的流动, 其流场存在流动变化剧烈和平缓的区域, 在流动变化剧烈的区域, 往往需要足够的网格点才能更好地捕捉到流场信息, 而均匀网格会使得网格数量过多, 这会增加计算量, 但网格数量过少又无法获得必要的流场信息, 使LBM的计算效率降低.为了解决上述问题, 用不同的网格结构, 以顶盖驱动的腔体内流、柱体绕流和翼型绕流为例, 探讨了提高LBM算法的计算效率和适用性问题.     

两个角区湍流场及其尾迹的实验研究

绕两个翼型－平面的角区流动及其尾迹的实验是在低湍流度风洞中完成的．在零攻角条件下，对翼型－平面的角区流场内诸参数，如翼型表面和平板面上的压力分布、绕翼型及尾迹区内的平均速度、脉动速度、湍动能、二阶关联量ｕ′ｖ′及ｕ′ｗ′进行了广泛的测量．通过对比，分析了这两种模型与平面所构成的角区及其尾迹区内的流动特性     

基于涡粘性模型的翼型湍流流场熵产率计算

利用有限体积法实现了基于非正交同位网格的SIMPLE算法。基于熵分析方法,采用涡粘性模型求解湍流熵产方程,系统研究了湍流模型对二维翼型绕流流场熵产率的影响。通过计算NACA0012翼型在来流雷诺数为2.88×106时,0°攻角~16.5°攻角范围内的翼型表面压力系数分布和升阻力特性,验证了算法及程序的正确性。结果表明,选择不同湍流模型时,翼型流场熵产的计算结果存在差异,湍流耗散是引起流场熵产的主要原因;翼型流场的熵产主要发生在翼型前缘区、壁面边界层和翼型尾流区域,流场熵产率与翼型阻力系数线性相关;当产生分离涡时,粘性耗散引起的熵产下降。     

基于Boltzmann模型方程的气体运动论统一算法研究

模型方程出发,研究确立含流态控制参数可描述不同流域气体流动特征的气体分子速度分布函数方程; 研究发展气体运动论离散速度坐标法, 借助非定常时间分裂数值计算方法和NND差分格式, 结合DSMC方法关于分子运动与碰撞去耦技术, 发展直接求解速度分布函数的气体运动论耦合迭代数值格式; 研制可用于物理空间各点宏观流动取矩的离散速度数值积分方法, 由此提出一套能有效模拟稀薄流到连续流不同流域气体流动问题统一算法. 通过对不同Knudsen数下一维激波内流动、二维圆柱、三维球体绕流数值计算表明, 计算结果与有关实验数据及其它途径研究结果(如DSMC模拟值、N-S数值解)吻合较好, 证实气体运动论统一算法求解各流域气体流动问题的可行性. 尝试将统一算法进行HPF并行化程序设计, 基于对球体绕流及类``神舟'返回舱外形绕流问题进行HPF初步并行试算, 显示出统一算法具有很好的并行可扩展性, 可望建立起新型的能有效模拟各流域飞行器绕流HPF并行算法研究方向. 通过将气体运动论统一算法推广应用于微槽道流动计算研究, 已初步发展起可靠模拟二维短微槽道流动数值算法; 通过对Couette流、Poiseuille流、压力驱动的二维短槽道流数值模拟, 证实该算法对微槽道气体流动问题具有较强的模拟能力, 可望发展起基于Boltzmann模型方程能可靠模拟MEMS微流动问题气体运动论数值计算方法研究途径.      

满足几何守恒律的WENO格式及其应用

对几何守恒律的来源进行了分析,发展了一种满足几何守恒律的WENO格式,并应用于翼型层流分离现象的数值模拟中。为消除网格质量影响,采用守恒型方法计算网格导数,并将标准的WENO格式分解为中心差分部分和数值耗散部分。算例计算结果表明,几何守恒律对高精度有限差分WENO格式计算结果具有重要影响,本文方法能够消除网格导数计算误差,保证来流保持性。将本文方法应用于SD7003翼型层流分离现象的数值模拟中,计算结果与文献中计算及试验数据吻合较好,同时能够精细捕捉小尺度流场结构,准确模拟翼型层流分离现象中的复杂流动过程。     

大型低速二元风洞侧壁边界层控制及对单段翼型实验结果的影响

本文简述了ＮＦ－３风洞二元实验段侧壁边界层吹除控制系统及具有吹气的模型实验方法，给出了不同吹气系数对风洞边界层的控制效果以及对相对厚度为７％的单段翼型实验结果的影响。初步实验研究结果表明，该控制系统能有效地改善风洞侧壁边界层的流动状态，减小侧壁干扰，改善翼型实验中的二元流动特性     

Turbulence modelling of the shear layers in the trailing edge region of a modern rear-loaded airfoil at incidence

After carefull analysis in a turbulent zero-pressure gradient flow, various simple algebraic turbulence models were applied to the almost separated flow on the upperside of an airfoil at incidence. The Johnson-King and Horton non-equilibrium (or rate equation) models give clearly improved results.     

基于涡粘性模型的翼型湍流流场熵产率计算

利用有限体积法实现了基于非正交同位网格的SIMPLE算法。基于熵分析方法,采用涡粘性模型求解湍流熵产方程,系统研究了湍流模型对二维翼型绕流流场熵产率的影响。通过计算NACA0012翼型在来流雷诺数为2.88×10⁶时,0°攻角~16.5°攻角范围内的翼型表面压力系数分布和升阻力特性,验证了算法及程序的正确性。结果表明,选择不同湍流模型时,翼型流场熵产的计算结果存在差异,湍流耗散是引起流场熵产的主要原因;翼型流场的熵产主要发生在翼型前缘区、壁面边界层和翼型尾流区域,流场熵产率与翼型阻力系数线性相关;当产生分离涡时,粘性耗散引起的熵产下降。     

Application of hybrid Cartesian grid and gridless approach to moving boundary flow problems

The paper presents a hybrid Cartesian grid and gridless approach to solve unsteady moving boundary flow problems. Unlike the Chimera clouds of points approach, the hybrid approach uses a Cartesian grid to cover most of the computational domain and a gridless method to calculate a relatively small region adjacent to the body surface, making use of the flexibility of the gridless method in handling surface grid with complicated geometry and the computational efficiency of the Cartesian grid. Four cases were conducted to examine the applicability, accuracy and robustness of the hybrid approach. Steady flows over a single NACA0012 airfoil and dual NACA0012 airfoils at different Mach numbers and angles of attack were simulated. Moreover, by implementing a dynamic hole cutting, node identification and information communication between the Cartesian grid and the gridless regions, unsteady flows over a pitching NACA0012 airfoil (small displacement) and two‐dimensional airfoil/store separation (large displacement) were performed. The computational results were found to agree well with earlier experimental data as well as computational results. Shock waves were accurately captured. The computational results show that the hybrid approach is of potential to solve the moving boundary flow problems. Copyright © 2013 John Wiley & Sons, Ltd.     

Mathematical modeling of the sound induced in a subsonic flow around cascades in relative motion

A method of mathematical modeling of the tonal sound induced by the unsteady aerodynamic interaction of two plane airfoil cascades in a subsonic flow and in uniform relative motion in the direction of their fronts is developed. The method is based on the numerical integration of the unsteady flow equations using a simplified model for the periodic system of edge wakes shed from the airfoils of the first (leading) cascade in the viscous flow and acting on the second (trailing) cascade. An analysis of the distinctive features of the flow under consideration demonstrates the efficiency of the model proposed.     

Experimental study of pitching and plunging airfoils at low Reynolds numbers

Measurements of the unsteady flow structure and force time history of pitching and plunging SD7003 and flat plate airfoils at low Reynolds numbers are presented. The airfoils were pitched and plunged in the effective angle of attack range of 2.4°–13.6° (shallow-stall kinematics) and ?6° to 22° (deep-stall kinematics). The shallow-stall kinematics results for the SD7003 airfoil show attached flow and laminar-to-turbulent transition at low effective angle of attack during the down stroke motion, while the flat plate model exhibits leading edge separation. Strong Re-number effects were found for the SD7003 airfoil which produced approximately 25 % increase in the peak lift coefficient at Re = 10,000 compared to higher Re flows. The flat plate airfoil showed reduced Re effects due to leading edge separation at the sharper leading edge, and the measured peak lift coefficient was higher than that predicted by unsteady potential flow theory. The deep-stall kinematics resulted in leading edge separation that led to formation of a large leading edge vortex (LEV) and a small trailing edge vortex (TEV) for both airfoils. The measured peak lift coefficient was significantly higher (~50 %) than that for the shallow-stall kinematics. The effect of airfoil shape on lift force was greater than the Re effect. Turbulence statistics were measured as a function of phase using ensemble averages. The results show anisotropic turbulence for the LEV and isotropic turbulence for the TEV. Comparison of unsteady potential flow theory with the experimental data showed better agreement by using the quasi-steady approximation, or setting C(k) = 1 in Theodorsen theory, for leading edge–separated flows.     

A novel method for automated grid generation of ice shapes for local‐flow analysis

Modelling a complex geometry, such as ice roughness, plays a key role for the computational flow analysis over rough surfaces. This paper presents two enhancement ideas in modelling roughness geometry for local flow analysis over an aerodynamic surface. The first enhancement is use of the leading‐edge region of an airfoil as a perturbation to the parabola surface. The reasons for using a parabola as the base geometry are: it resembles the airfoil leading edge in the vicinity of its apex and it allows the use of a lower apparent Reynolds number. The second enhancement makes use of the Fourier analysis for modelling complex ice roughness on the leading edge of airfoils. This method of modelling provides an analytical expression, which describes the roughness geometry and the corresponding derivatives. The factors affecting the performance of the Fourier analysis were also investigated. It was shown that the number of sine–cosine terms and the number of control points are of importance. Finally, these enhancements are incorporated into an automated grid generation method over the airfoil ice accretion surface. The validations for both enhancements demonstrate that they can improve the current capability of grid generation and computational flow field analysis around airfoils with ice roughness. Copyright © 2004 John Wiley & Sons, Ltd.     

Finite element computation of turbulent flow past a multi-element airfoil

Flow past multi-element airfoil is studied via two-dimensional numerical simulations. The incompressible Reynolds averaged Navier–Stokes equations, in primitive variables, are solved using a stabilized finite element formulation. The Spalart–Allmaras and Baldwin–Lomax models are employed for turbulence closure. The implementation of the Spalart–Allmaras model is verified by computing flow over a flat plate with a specified trip location. Good agreement is seen between the results obtained with the two models for flow past a NACA 0012 airfoil at 5° angle of attack. Results for the multi-element airfoil, with the two turbulence models, are compared with experiments for various angles of attack. In general, the pressure distribution, from both the models matches quite well with the experimental results. However, at larger angles of attack, the computational results overpredict the suction peak on the slat. The velocity profiles from the Baldwin–Lomax model are, in general, more diffused compared to those from the Spalart–Allmaras model. The agreement between the computed and experimental results is not too good in the flap region for large angles of attack. Both the models are unable to predict the stall; the flow remains attached even for relatively large angles of attack. Consequently, the lift coefficient is over predicted at large α by the computations. Overall, compared to the Baldwin–Lomax model, the predictions from the Spalart–Allmaras model are closer to experimental measurements.     

The physics of modeling unsteady flaps with gaps

Research indicates that active control concepts have promise in mitigating numerous adverse phenomena associated with the aeromechanics of lifting surfaces. These techniques are being applied to delay stall of fixed wing aircraft, as well as to eliminate or mitigate vibratory loads, blade–vortex interaction, and dynamic stall of the flow about rotorcraft and wind turbine blades. These phenomena are nonlinear and unsteady for dynamic systems, which add yet another layer of complexity on the physics of the flow. While a plethora of different active control techniques is being explored, the use of trailing edge flaps appears to be one of the more viable and cost-effective concepts. Static multi-element airfoils and wings have been analyzed computationally, but little exists on the ability to model these when the airfoil and flap are dynamic. The costs associated with modeling the gap between the airfoil and flap have led to approximations where the flap is modeled only as a morphed tip of the airfoil (no gap). Using a hybrid Reynolds-Averaged Navier–Stokes/Large-Eddy-Simulation turbulence technique, an oscillating flapped airfoil has been studied to determine the influence of modeling the gap on the performance and acoustic signature of the airfoil. Results are compared with the experimental data to confirm the validity of the computational approach. Both attached and separated (dynamic stall) oscillating flows are examined. The physics within the gap are found to be important for the airfoil performance when stall is encountered, as well as when acoustic signatures are required.     

A study of long separation bubble on thick airfoils and its consequent effects

A parametric study has been performed to analyse the flow around the thick-symmetric NACA 0021 airfoil in order to better understand the characteristics and effects of long separation bubbles (LoSBs) that exist on such airfoils at low Reynolds numbers and turbulence intensities. In the article, the prediction capabilities of two recently-developed transition models, the correlation-based γ–Re_θ model and the laminar-kinetic-energy-based κ–κ_L–ω model are assessed. Two-dimensional steady-state simulations indicated that the κ–κ_L–ω model predicted the separation and reattachment process accurately when compared with published experimental work. The model was then used to study the attributes and the effects of LoSBs as a function of the angle of attack, freestream turbulence intensity and Reynolds number. It was observed that LoSBs considerably degrade the aerodynamic performance of airfoils and lead to abrupt stall behaviour. It is, furthermore, illustrated that the presence of the LoSB leads to an induced camber effect on the airfoil that increases as the airfoil angle of attack increases due to the upstream migration of the bubble. An increase in the Reynolds number or turbulence levels leads to a reduction in the bubble extent, considerably improving the airfoil performance and leading to a progressive trailing-edge stall.