基于网格框架的非结构附面层网格生成技术

附面层网格质量是确保计算流体力学粘性计算精度的关键技术环节.本文针对复杂外形提出了全局一致的高质量附面层网格构造算法,该方法基于针对特征的表面网格区域分解技术,利用表面网格分片后的边界线及其法向量构造最终附面层网格的轮廓框架线,并通过径向基函数及线性插值算法生成完整的附面层网格.通过典型算例分析可以看出,该方法生成的非结构附面层网格精度和全局一致性较高,且能够有效避免复杂外形附面层网格局部及全局交叉现象.     

高雷诺数下求解NS方程的无网格算法

提出了一种适合高雷诺数NS方程求解的隐式无网格算法。针对高雷诺数粘性流动的特点,在附面层内的粘性影响区域采用法向层次推进布点的方法形成离散点云,在附面层外的计算区域内实行填充式布点的方法形成离散点云。根据附面层内外点云的不同构造特点,推导出运用格林公式和最小二乘曲面拟合方法求取空间导数的统一形式,在此基础上运用AUSM _up格式求得数值通量,并引入BL湍流模型对雷诺平均NS方程的湍流应力项进行封闭。时间推进格式方面,采用了计算效率较高的隐式高斯-赛德尔迭代算法。为了验证本文方法的计算精度和鲁棒性,对NACA0012翼型低速流动、RAE2822翼型跨音速绕流和二维圆柱的分离流动进行了数值模拟。     

结合自适应网格的基于特征线方程的分离算法

提出一种新的网格自适应方法:在需要加密的网格单元中心加入新结点,并对加密后的相邻三角形网格单元进行公共边变换, 构成新的网格单元. 与传统的在网格边界中点加入新节点的自适应方法相比,新方法可以更加灵活地控制网格密度,加密后的网格继承原先的网格质量不发生畸变,并且算法编程简便,容易实现. 将自适应网格生成方法和基于特征线方程的分离算法相结合,对空腔内不可压缩黏性流动进行了计算. 在特征线方向上进行时间步离散,动量方程求解过程中采用非增量型分离算法. 计算中,把求解变量梯度值作为判定准则,在变化剧烈的区域进行网格局部加密. 计算结果表明该组合算法有很好的计算精度,并有效减少了计算时间和存储量.      

计算含大位移动边界非定常流动的无网格算法

研究了可以计算处理包含大位移动边界非定常流动的无网格算法.创建了无网格方法中由于大位移边界运动造成的不合要求点的判断标准,采用高效的填充方法实现点云重构,运用线性插值方法得到新点参数,实现了局部重构处理大位移动点.流场计算方面,在计算域自动布点基础上,采用曲线逼近计算导数及HLLC格式计算数值通量,发展了求解基于无网格的ALE方程组的算法.最后,对活塞问题进行了模拟,结果与解析解相吻合,验证了算法的准确性;另外,计算比较了气流流过静止圆柱以及圆柱在静止流场运动流场,结果表明方法是成功的.     

福建海岸带地质环境背景与海水入侵的形成条件探讨

提出一种新的网格自适应方法：在需要加密的网格单元中心加入新结点,并对加密后的相邻 三角形网格单元进行公共边变换, 构成新的网格单元. 与传统的在网格边界中点加入新节点的自 适应方法相比,新方法可以更加灵活地控制网格密度,加密后的网格继承原先的网格质量不 发生畸变,并且算法编程简便,容易实现. 将自适应网格生成方法和基于特征线方程的分离 算法相结合,对空腔内不可压缩黏性流动进行了计算. 在特征线方向上进行时间步离散,动 量方程求解过程中采用非增量型分离算法. 计算中,把求解变量梯度值作为判定准则,在 变化剧烈的区域进行网格局部加密. 计算结果表明该组合算法有很好的计算精度,并有效减 少了计算时间和存储量.     

计算含动边界非定常流动的无网格算法

在无网格算法中考虑了含动边界的流动问题,研究了可以计算处理包含一定位移及扭转动边界非定常流动的算法.创建了无网格算法的动点法则,并引入抗扭方法对弹簧方法进行改进来处理离散点运动,提高了方法的可用度及精度.发展了求解基于无网格的ALE方程组的算法,在点云离散的基础上采用曲面逼近计算空间导数及HLLC格式计算数值通量,运用四步龙格-库塔法进行时间推进.在跨、超音速条件下,计算模拟了典型翼型简谐振动流场,计算结果与实验结果及文献对比吻合,验证了该算法的正确性.     

翼身组合体气动力计算研究

采用求解Euler方程结合附面层修正的方法在结构网格上对翼身组合体跨音速流场进行了数值模拟.附面层方程的求解应用Whitfield提出的动量积分方程和平均流动能积分方程,为了保持Euler方程求解过程中计算网格的固定性,用加在物面上的溢出速度来模拟附面层效应.针对传统的近场方法计算阻力,计算精度较低、误差较大并且不能给出各阻力分量值的缺点,将基于动量定理的远场方法用于飞机的阻力估算,采用远场法将阻力分解为:粘性阻力,激波阻力,诱导阻力,并对各个分量分别进行了求解,将计算结果与近场法以及风洞实验值做了比较.以DLR-F4翼身组合体为考核算例,对所述方法进行了验证,结果显示远场法的计算结果与风洞实验值吻合的很好.     

基于无网格一图论方法的岩体边坡稳定性分析研究

将无网格伽辽金法应用于岩体边坡稳定性分析,发展了基于无网格模型和有向加权图Bellman—Ford最短路径搜索算法相结合的无网格-图论边坡滑移面搜索方法,以搜寻节理岩体边坡失稳时的临界滑移面并得出其相应的安全系数。区别于传统的滑移面搜索算法,本文方法无需假定滑移面形状,更适用于具有复杂滑移线形状的节理岩体边坡的稳定性分析与计算,具有稳定和高效的算法特点。文中详细论述了无网格-图论最短路径算法的理论、方法和程序实现,并通过算例说明该方法在岩体边坡稳定性分析中的适用性。     

基于无网格—图论方法的岩体边坡稳定性分析研究

将无网格伽辽金法应用于岩体边坡稳定性分析,发展了基于无网格模型和有向加权图Bellman-Ford最短路径搜索算法相结合的无网格-图论边坡滑移面搜索方法,以搜寻节理岩体边坡失稳时的临界滑移面并得出其相应的安全系数.区别于传统的滑移面搜索算法,本文方法无需假定滑移面形状,更适用于具有复杂滑移线形状的节理岩体边坡的稳定性分析与计算,具有稳定和高效的算法特点.文中详细论述了无网格-图论最短路径算法的理论、方法和程序实现,并通过算例说明该方法在岩体边坡稳定性分析中的适用性.     

基于AFT-Delaunay的二维解耦并行网格生成算法

面向平面任意几何区域网格生成,提出了一种将波前法AFT(Advancing Front Technique)与Delaunay法相结合的解耦并行网格生成算法。算法主要思想是沿着求解几何区域惯性轴,采用扩展的AFT-Delaunay算法生成高质量三角形网格墙,递归地将几何区域动态划分成多个彼此解耦的子区域;采用OpenMP多线程并行技术,将子区域分配给多个CPU并行生成子区域网格;子区域内部的网格生成复用AFT-Delaunay算法,保证了生成网格的质量、效率和一致性要求。本算法优先生成几何边界与交界面网格,有利于提高有限元计算精度;各个子区域的网格生成彼此完全解耦,因此并行网格生成过程无需通信。该方法克服了并行交界面网格质量恶化难题,且具有良好的并行加速比,能够全自动、高效率地并行生成高质量的三角网格。     

A vertex-centred finite volume method with shock detection

This paper introduces a vertex-centred finite volume method for compressible viscous flow incorporating a new shock detection procedure. The discretization is designed to be robust and accurate on the highly stretched and curved meshes necessary for resolving turbulent boundary layers around the leading edge of an aerofoil. Details of the method are described for two-dimensional problems and the natural extension of three-dimensional multiblock meshes is discussed. The shock detection procedure is used to limit the range of the shock-capturing dissipation specifically to regions containing shocks. For transonic turbulent flow this is shown to improve the boundary layer representation significantly.     

无网格算法在多段翼型流动计算中的应用

研究了一种求解欧拉方程的无网格算法，发展出了一套布点及点云自动生成的方法；在点云离散的基础上，采用最小二乘法求解矛盾方程的方法来求取空间导数，进而获得数值通量；采用四步龙格-库塔方法进行时间推进，并引入当地时间步长和残值光顺等加速收敛措施。通过对NA-CA0012翼型的跨音速流动和多段翼型复杂绕流的数值模拟，验证了上述无网格算法的正确性和实用性。     

Prediction of oscillating thick cambered aerofoil aerodynamics by a locally analytic method

A complete first-order model and locally analytic solution method are developed to analyse the effects of mean flow incidence and aerofoil camber and thickness on the incompressible aerodynamics of an oscillating aerofoil. This method incorporates analytic solutions, with the discrete algebraic equations which represent the differential flow field equations obtained from analytic solutions in individual grid elements. The velocity potential is separated into steady and unsteady harmonic parts, with the unsteady potential further decomposed into circulatory and non-circulatory components. These velocity potentials are individually described by Laplace equations. The steady velocity potential is independent of the unsteady flow field. However, the unsteady flow is coupled to the steady flow field through the boundary conditions on the oscillating aerofoil. A body-fitted computational grid is then utilized. Solutions for both the steady and the coupled unsteady flow fields are obtained by a locally analytic numerical method in which locally analytic solutions in individual grid elements are determined. The complete flow field solution is obtained by assembling these locally analytic solutions. This model and solution method are shown to accurately predict the Theodorsen oscillating flat plate classical solution. Locally analytic solutions for a series of Joukowski aerofoils demonstrate the strong coupling between the aerofoil unsteady and steady flow fields, i.e. the strong dependence of the oscillating aerofoil aerodynamics on the steady flow effects of mean flow incidence angle and aerofoil camber and thickness.     

ANALYSIS/DESIGN CALCULATIONS OF TRANSONIC FLOWS

SUMMARY

Analysis/design calculations of transonic flow are discussed and several improvements are made. The nonisentropic potential method is used to calculate the inviscid transonic flow analysis problem instead of the traditional potential method. An inverse integral 3D boundary layer method is used to calculate the boundary layer in the viscous transonic flow analysis problem. The viscous/inviscid interaction calculations are carried out by a semi-inverse coupling scheme. In design problem calculations, an improved residual-correction method is used. Three individual methods are combined in a global algorithm and computing code. The improvements speed up the convergence, increase applicability and computational efficiency. Some numerical results are given to illustrate that the present method provides an effective engineering tool of high accuracy and efficiency in three dimensional transonic analysis and design situations.     

The influence of acceleration and deceleration on shock wave movement on and around aerofoils in transonic flight

This paper examines the shock wave dynamics of a biconvex aerofoil in transonic flight during acceleration and retardation. The aerofoil has a cord length of 1 m and air at infinity is at 101.325 kPa and 300 K. Using Fluent as the CFD software, constant velocity (steady state) simulations were conducted at transonic Mach numbers. The aerofoil was then accelerated at 1041m/s² (106 g), starting at Mach 0.1, and decelerated at −1041m/s², starting at Mach 1.6, through the same range of Mach numbers using time-dependent (unsteady) simulations. Significant differences were found in the transonic region between the steady and the unsteady aerodynamic forces. Analysis of the flow field in this region showed that acceleration-dependent variations in the position of the shock wave on the surfaces of the aerofoil were the main reason for this. As very high accelerations were used in order to emphasize differences, which do not have many practical applications, simulations using accelerations lower than 9 g were also conducted in order to confirm the results. The acceleration-dependent behaviour of other shock waves around the aerofoil, such as the bow shock in front of the aerofoil and the trailing wave were also examined. The trailing wave followed behind the aerofoil changing position with different accelerations at the same Mach number.      

Transonic shock oscillations on NACA0012 aerofoil

The mechanism of the origin of shock oscillations on NACA0012 aerofoils is investigated using a moving grid thin layer Navier Stokes code. The method used to understand the mechanism is to initiate the shock oscillations on an aerofoil by moving the aerofoil from a regime of steady transonic flow into a regime of periodic flow by a change in airflow incidence. The results indicate that the shock induced bubble plays a leading role in the origin of shock oscillations and the trailing edge has an affect on its amplitude. Received 1 April 1997 / Accepted 1 December 1997     

NUMERICAL UNCERTAINTIES IN TRANSONIC FLOW CALCULATIONS FOR AEROFOILS WITH BLUNT TRAILING EDGES

Numerical uncertainties are quantified for calculations of transonic flow around a divergent trailing edge (DTE) supercritical aerofoil. The Reynolds-averaged Navier–Stokes equations are solved using a linearized block implicit solution procedure and mixing-length turbulence model. This procedure has reproduced measurements around supercritical aerofoils with blunt trailing edges that have shock, boundary layer and separated regions. The present effort quantifies numerical uncertainty in these calculations using grid convergence indices which are calculated from aerodynamic coefficients, shock location, dimensions of the recirculating region in the wake of the blunt trailing edge and distributions of surface pressure coefficients. The grid convergence index is almost uniform around the aerofoil, except in the shock region and at the point where turbulence transition was fixed. The grid convergence index indicates good convergence for lift but only fair convergence for moment and drag and also confirms that drag calculations are more sensitive to numerical error. © 1997 by John Wiley and Sons, Ltd.     

AUFS 格式在无网格方法中的应用简

将计算量小,激波分辨率高的AUFS (artificially upstream flux vector splitting) 格式应用于无网格方法. 所发展算法基于多项式基函数最小二乘无网格方法,采用线性基函数曲面拟合及AUFS 格式计算各离散点的空间导数,应用四阶Runge-Kutta 法进行时间显式推进. 为验证算法健壮性、精度以及计算效率,对Riemann 问题、超音速平面流动,以及不同攻角NACA0012 翼型跨音速流场进行了数值模拟,其结果同采用HLLC (Harten-Lax-van Leer-contact) 格式的无网格方法以及文献报道结果吻合较好,并且计算量较形式简单HLLC 格式减少约15%.     

A multi-physics computational model of fuel sloshing effects on aeroelastic behaviour

A multi-physics computational method is presented to model the effect of internally and externally-carried fuel on aeroelastic behaviour of a pitch–plunge aerofoil model through the transonic regime. The model comprises three strongly coupled solvers: a compressible finite-volume Euler code for the external flow, a two-degree of freedom spring model and a smoothed particle hydrodynamics solver for the fuel. The smoothed particle hydrodynamics technique was selected as this brings the benefit that nonlinear behaviour such as wave breaking and tank wall impacts may be included. Coupling is accomplished using an iterative method with subcycling of the fuel solver to resolve the differing timestep requirements. Results from the fuel-structural system are validated experimentally, and internally and externally-carried fuel is considered using time marching analysis. Results show that the influence of the fuel, ignoring the added mass effect, is to raise the flutter boundary at transonic speeds, but that this effect is less pronounced at lower Mach numbers. The stability boundary crossing is also found to be less abrupt when the effect of fuel is included and limit cycles often appear. An external fuel tank is seen to exhibit a lower stability boundary, while the response shows a beating effect symptomatic of two similar frequency components, potentially due to interaction between vertical and horizontal fuel motion.     

Existence and Stability of Multidimensional Transonic Flows through an Infinite Nozzle of Arbitrary Cross-Sections

We establish the existence and stability of multidimensional steady transonic flows with transonic shocks through an infinite nozzle of arbitrary cross-sections, including a slowly varying de Laval nozzle. The transonic flow is governed by the inviscid potential flow equation with supersonic upstream flow at the entrance, uniform subsonic downstream flow at the exit at infinity, and the slip boundary condition on the nozzle boundary. Our results indicate that, if the supersonic upstream flow at the entrance is sufficiently close to a uniform flow, there exists a solution that consists of a C ^1,α subsonic flow in the unbounded downstream region, converging to a uniform velocity state at infinity, and a C ^1,α multidimensional transonic shock separating the subsonic flow from the supersonic upstream flow; the uniform velocity state at the exit at infinity in the downstream direction is uniquely determined by the supersonic upstream flow; and the shock is orthogonal to the nozzle boundary at every point of their intersection. In order to construct such a transonic flow, we reformulate the multidimensional transonic nozzle problem into a free boundary problem for the subsonic phase, in which the equation is elliptic and the free boundary is a transonic shock. The free boundary conditions are determined by the Rankine–Hugoniot conditions along the shock. We further develop a nonlinear iteration approach and employ its advantages to deal with such a free boundary problem in the unbounded domain. We also prove that the transonic flow with a transonic shock is unique and stable with respect to the nozzle boundary and the smooth supersonic upstream flow at the entrance.