基于涡量矩理论的绕振荡水翼涡动力学分析

文章采用标准k-ω SST湍流模型和动网格技术, 实现了绕俯仰振荡NACA66水翼非定常流动结构与水动力特性的数值模拟, 并基于有限域涡量矩理论定量表征了局部旋涡结构对水翼动力特性的影响. 研究结果表明: 在水翼升程阶段, 当攻角较小时, 层流向湍流的转捩点由水翼尾缘向前缘移动; 在较大攻角时, 顺时针尾缘涡?TEV在水翼吸力面上生成并向前缘发展, 同时与吸力面上的顺时针前缘涡?LEV融合发展为附着在整个吸力面上的新前缘涡?LEV, 新的?LEV与逆时针尾缘涡+TEV相互作用直至完全脱落, 直接导致了水翼的动力失速, 在回程阶段, 绕振荡水翼的流场结构逐渐由湍流转变为层流. 基于有限域涡量矩理论的定量分析发现, 有限域内附着的?LEV和?TEV提供正升力, 当?LEV发展覆盖整个吸力面时对升力的贡献最大, 占总升力近50%, 而+TEV提供负升力. 同时发现, 有限域内各旋涡内部的不同区域提供的升力有正有负; 而逸出有限域的旋涡内部不同区域提供的升力方向均保持一致, 其中顺时针涡提供正升力, 而逆时针涡提供负升力. 在失速阶段, 域外旋涡整体对升力贡献较小且存在小幅波动, 体现了流动的非定常性.      

等速上仰翼型分离流动结构的研究

本文通过在翼型上游和翼表面边界层内放置产生氢气泡的铂丝的方法,清楚地显示了上仰翼型分离剪切层的结构。揭示了在不同的翼型转动角速度范围内,存在三种分离流结构。研究了失速涡,剪切涡及起动涡随时间的演变,它们之间的相互作用和转动角速度等参数的影响,分离剪切层的流动显示结果,结合翼型上气动力与流场中涡量矩的关系的理论,定性地解释了上仰翼型产生非定常高升力的原因。     

NUMERICAL STUDY ON DYNAMIC STALL FLOW CONTROL FOR WIND TURBINE AIRFOIL USING PLASMA ACTUATOR 1)

针对动态失速引起的风力机翼型气动性能恶化的问题,本文基于动网格和滑移网格技术, 开展了大涡模拟数值计算研究,探索了非定常脉冲等离子体的动态流动控制机理. 结果表明,等离子体气动激励能够有效控制翼型动态失速, 改善平均和瞬态气动力,减小力矩负峰值和迟滞环面积. 压力分布在等离子体施加范围内出现了负压"凸起",上翼面吸力峰值明显增大.脉冲频率和占空比这两个非定常控制参数对流动控制影响显著,无因次脉冲频率为1.5时等离子体控制效果较好,占空比为0.8时即可接近连续工作模式下的气动收益. 翼型深失速状态,等离子体促使流动分离位置明显向后缘移动, 抵抗了大尺度动态失速涡的发生,分离涡结构破碎耗散、重新附着, 涡流影响范围减小; 浅失速状态,等离子体激励具有较强的剪切层操纵能力, 诱导了翼型边界层提前转捩,促进了与主流的动量掺混. 等离子体气动激励诱导出前缘附近贴体翼面"涡簇",起到了虚拟气动外形的作用.不同尺度、频域的动态涡结构与等离子体气动激励的非线性、强耦合作用导致了气动力/力矩的谐波振荡.     

风力机翼型动态失速等离子体流动控制数值研究

针对动态失速引起的风力机翼型气动性能恶化的问题,本文基于动网格和滑移网格技术, 开展了大涡模拟数值计算研究,探索了非定常脉冲等离子体的动态流动控制机理. 结果表明,等离子体气动激励能够有效控制翼型动态失速, 改善平均和瞬态气动力,减小力矩负峰值和迟滞环面积. 压力分布在等离子体施加范围内出现了负压"凸起",上翼面吸力峰值明显增大.脉冲频率和占空比这两个非定常控制参数对流动控制影响显著,无因次脉冲频率为1.5时等离子体控制效果较好,占空比为0.8时即可接近连续工作模式下的气动收益. 翼型深失速状态,等离子体促使流动分离位置明显向后缘移动, 抵抗了大尺度动态失速涡的发生,分离涡结构破碎耗散、重新附着, 涡流影响范围减小; 浅失速状态,等离子体激励具有较强的剪切层操纵能力, 诱导了翼型边界层提前转捩,促进了与主流的动量掺混. 等离子体气动激励诱导出前缘附近贴体翼面"涡簇",起到了虚拟气动外形的作用.不同尺度、频域的动态涡结构与等离子体气动激励的非线性、强耦合作用导致了气动力/力矩的谐波振荡.      

等速上仰翼型动态失速现象研究

翼型大迎角绕流的静态失速将造成升力突降和气动性能急剧恶化，但利用非定常运动所产生 的动态失速效应，可以大大地延缓气流分离和失速现象的发生. 采用Rogers发 展的双时间步Roe格式，求解拟压缩性修正不可压N-S方程. 数值模拟了低雷诺数 ($Re=4.8 \times 10^{4}$)条件下NACA0015翼型作等速上仰($\alpha =0^{\circ} \sim 60^{\circ}$)的动态失速过程，同Walker的试验结果比 较，验证了计算结果的正确性. 研究了该过程中主涡、二次涡和三次涡的发展，升 力系数随攻角变化，以及不同上仰速度对动态失速效应所造成的影响.     

旋翼翼型动态失速流场特性PIV试验研究及L-B模型修正

为测量翼型动态失速的非定常涡流场特性,采用3D-PIV 技术,对典型直升机旋翼翼型SC1095 的动态失速流场特性进行测量,发现涡在不同位置处的输运速度不同:位于翼型表面的涡的无量纲速度为0.39,位于尾迹区的涡的无量纲速度为0.55. 利用前缘涡输运速度变化这一特征,改进了经典的翼型动态失速利什曼-贝多斯（Leishman-Beddoes,L-B）模型,将该模型中固定的涡时间常数修正为可以随涡位置变化的时变函数,修正后的模型计算得到翼型法向力峰值相对原L-B 模型提升5%,力矩系数负峰值相对原L-B 模型提升13%,与试验值相比更加吻合,表明修正后的翼型动态失速模型更好地体现了翼型前缘涡的物理特征.      

PIV EXPERIMENTS ON FLOWFIELD CHARACTERISTICS OF ROTOR AIRFOIL DYNAMIC STALL AND MODIFICATIONS OF L-B MODEL

为测量翼型动态失速的非定常涡流场特性,采用3D-PIV 技术,对典型直升机旋翼翼型SC1095 的动态失速流场特性进行测量,发现涡在不同位置处的输运速度不同：位于翼型表面的涡的无量纲速度为0.39,位于尾迹区的涡的无量纲速度为0.55. 利用前缘涡输运速度变化这一特征,改进了经典的翼型动态失速利什曼-贝多斯（Leishman-Beddoes,L-B）模型,将该模型中固定的涡时间常数修正为可以随涡位置变化的时变函数,修正后的模型计算得到翼型法向力峰值相对原L-B 模型提升5%,力矩系数负峰值相对原L-B 模型提升13%,与试验值相比更加吻合,表明修正后的翼型动态失速模型更好地体现了翼型前缘涡的物理特征.     

动失速型非定常分离涡结构的控制

本文对动失速型非定常分离涡结构的控制方法，在低速风洞中应用相平均测压技术进行了实验研究。在二元平板模型中部安装一作俯仰振荡的扰流板产生动失速型分离涡，在其上游安装另一用作控制的小扰流板。实验结果表明，应用前置的振荡小扰流板可影响并改变动失速分离涡的强度和对流特性。在最有利的控制相位下，涡吸力峰可降低４８％，涡对流时间可以推迟０．１９周期。对于间歇式振荡扰流板，采用相位提前控制方式比相位滞后控制方式更有效。     

波涡相互作用研究的某些进展(Ⅰ)

波涡相互作用是当代非定常非线性流体力学的一个重要领域.本文评述了迄今人们对微扰纵波(声波,不包括激波)和横波(涡波)与层状和轴状涡相互作用的研究进展,包括涡声理论、锐缘及光滑表面涡层分离以及自由涡层的2维感受性和非定常Kutta条件,波产生时均涡流的非线性Eulcr整流和广义Lagrange整流,波涡共振等课题的进展.还介绍了用非定常激发实施涡控制这一重要应用领域的一些新进展.      

模型昆虫翼作非定常运动时的气动力特性

基于Ｎａｖｉｅｒ－Ｓｔｏｋｅｓ方程的数值解,研究了一模型昆虫翼在小雷诺数（Ｒｅ＝１００）下作非定常运动时的气动力特性,这些运动包括：翼启动后的常速转动,快速加、减速转动,常速转动中快速上仰（模拟昆虫翼的上挥或下拍、翻转等运动）有如下结果：在小雷诺数下,模型昆虫翼以大攻角（α＝３５°）作常速转动运动时,由于失速涡不脱落,可产生较大的升为系数。其机理是：翼转动时,翼尖附近（该处线速度大）上翼面压强比翼根附近（该处线速度小）的小得多,因而存在展向压强梯度,同时存在着沿展向的离心力,此展向压强梯度和离心力导致的展向流动在失速涡的轴向方向,其可避免失速涡脱落。模型昆虫翼在快速加、减速转动和快速上仰运动中,虽然雷诺数小,但由于在短时间内产生了大涡量,也可产生十分大的气动力,例如在快速上仰运动中,升力系数可大于１０。     

An experimental study of the effects of pitch-pivot-point location on the propulsion performance of a pitching airfoil

An experimental investigation was conducted to characterize the evolution of the unsteady vortex structures in the wake of a pitching airfoil with the pitch-pivot-point moving from 0.16C to 0.52C (C is the chord length of the airfoil). The experimental study was conducted in a low-speed wind tunnel with a symmetric NACA0012 airfoil model in pitching motion under different pitching kinematics (i.e., reduced frequency k=3.8–13.2). A high-resolution particle image velocimetry (PIV) system was used to conduct detailed flow field measurements to quantify the characteristics of the wake flow and the resultant propulsion performance of the pitching airfoil. Besides conducting “free-run” PIV measurements to determine the ensemble-averaged velocity distributions in the wake flow, “phase-locked” PIV measurements were also performed to elucidate further details about the behavior of the unsteady vortex structures. Both the vorticity–moment theorem and the integral momentum theorem were used to evaluate the effects of the pitch-pivot-point location on the propulsion performance of the pitching airfoil. It was found that the pitch-pivot-point would affect the evolution of the unsteady wake vortices and resultant propulsion performance of the pitching airfoil greatly. Moving the pitch-pivot-point of the pitching airfoil can be considered as adding a plunging motion to the original pitching motion. With the pitch-pivot-point moving forward (or backward), the added plunging motion would make the airfoil trailing edge moving in the same (or opposite) direction as of the original pitching motion, which resulted in the generated wake vortices and resultant thrust enhanced (or weakened) by the added plunging motion.     

Investigation of three-dimensional dynamic stall on an airfoil using fast-response pressure-sensitive paint

Dynamic stall on a pitching OA209 airfoil in a wind tunnel is investigated at Mach 0.3 and 0.5 using high-speed pressure-sensitive paint (PSP) and pressure measurements. At Mach 0.3, the dynamic stall vortex was observed to propagate faster at the airfoil midline than at the wind-tunnel wall, resulting in a “bowed” vortex shape. At Mach 0.5, shock-induced stall was observed, with initial separation under the shock foot and subsequent expansion of the separated region upstream, downstream and along the breadth of the airfoil. No dynamic stall vortex could be observed at Mach 0.5. The investigation of flow control by blowing showed the potential advantages of PSP over pressure transducers for a complex three-dimensional flow.     

Influence of pitching angle of incidence on the dynamic stall behavior of a symmetric airfoil

This study presents the influence of pitch angle of an airfoil on its near-field vortex structure as well as the aerodynamic loads during a dynamic stall process. Dynamic stall behavior in a sinusoidally pitching airfoil is usually analyzed at low to medium reduced frequencies and with the maximum angle of attack of the airfoil not exceeding 25°. In this work, we study dynamic stall of a symmetric airfoil at medium to high reduced frequencies even as the maximum angle of attack goes from 25° to 45°. The evolution and growth of the laminar separation bubble, also known as a dynamic stall vortex, at the leading edge and the trailing edge are studied as the pitch cycle goes from the minimum to the maximum angle of attack. The effect of reduced frequencies on the vortex structure as well as the aerodynamic load coefficients is investigated. The reduced frequency is shown to be a bifurcation parameter triggering period doubling behavior. However, the bifurcation pattern is dependent on the variation of the pitch angle of incidence of the airfoil.     

Particle image velocimetry investigation of flow over unsteady airfoil with trailing-edge strip

The flow over a flapped NACA 0012 airfoil, oscillated slightly through the static-stall angle, was investigated by using particle image velocimetry, and was supplemented by surface pressure and dynamic-load measurements. A significant increase in the dynamic lift force and nose-down pitching moment was observed. The most pronounced flow phenomenon was the formation and detachment of an energetic leading-edge vortex compared to the no-flapped airfoil. The details of the underlying physical mechanisms responsible for the various light-stall flow processes were provided via the instantaneous velocity and vorticity fields measurements. In contrast to the Gurney flap, the inverted trailing-edge strip led to an improved negative damping while a reduced lift force. The addition of an inverted strip always led to the appearance of a Karman-type vortex shedding street immediately downstream of the strip over the entire oscillation cycle.     

Flow field characteristics study of a flapping airfoil using computational fluid dynamics

The flow field of a flapping airfoil in Low Reynolds Number (LRN) flow regime is associated with complex nonlinear vortex shedding and viscous phenomena. The respective fluid dynamics of such a flow is investigated here through Computational Fluid Dynamics (CFD) based on the Finite Volume Method (FVM). The governing equations are the unsteady, incompressible two-dimensional Navier-Stokes (N-S) equations. The airfoil is a thin ellipsoidal geometry performing a modified figure-of-eight-like flapping pattern. The flow field and vortical patterns around the airfoil are examined in detail, and the effects of several unsteady flow and system parameters on the flow characteristics are explored. The investigated parameters are the amplitude of pitching oscillations, phase angle between pitching and plunging motions, mean angle of attack, Reynolds number (Re), Strouhal number (St) based on the translational amplitudes of oscillations, and the pitching axis location (x/c). It is shown that these parameters change the instantaneous force coefficients quantitatively and qualitatively. It is also observed that the strength, interaction, and convection of the vortical structures surrounding the airfoil are significantly affected by the variations of these parameters.     

BIFURCATIONS OF AIRFOIL IN INCOMPRESSIBLE FLOW

Bifurcations of an airfoil with nonlinear pitching stiffness in incompressible flow are investigated. The pitching spring is regarded as a spring with cubic stiffness. The motion equations of the airfoil are written as the four dimensional one order differential equations. Taking air speed and the linear part of pitching stiffness as the parameters, the analytic solutions of the critical boundaries of pitchfork bifurcations and Hopf bifurcations are obtained in 2 dimensional parameter plane. The stabilities of the equilibrium points and the limit cycles in different regions of 2 dimensional parameter plane are analyzed. By means of harmonic balance method, the approximate critical boundaries of 2-multiple semi-stable limit cycle bifurcations are obtained, and the bifurcation points of supercritical or subcritical Hopf bifurcation are found. Some numerical simulation results are given.     

绕振荡水翼流动及其转捩特性的数值计算研究简

通过对比标准k-ω SST 湍流模型和基于标准k-ω SST 湍流模型修正的γ-Re_θ 转捩湍流模型对绕振荡NACA66 水翼流动的数值计算结果与实验结果,对水翼振荡过程的水动力特性和流场结构变化进行了分析研究. 结果表明：与标准k-ω SST 湍流模型的数值计算结果相比,基于标准k-ω SST 湍流模型修正的γ-Re_θ 转捩湍流模型能有效预测绕振荡翼型流场结构和水动力特性,捕捉流场边界层发生的流动分离和转捩现象;绕振荡水翼的流动过程可分为5 个特征阶段,当来流攻角较小时,在水翼前缘发生层流向湍流的转捩现象,水翼动力特征曲线出现变化拐点;随着来流攻角的增大,顺时针尾缘涡逐渐形成并向水翼前缘发展;当攻角较大时,前缘涡分离导致动力失速,水翼的动力特征曲线出现大幅波动;水翼处于顺时针向下旋转阶段,绕水翼的流动状态逐渐由湍流过渡为层流.     

绕振荡水翼流动及其转捩特性的数值计算研究

通过对比标准k-ω SST 湍流模型和基于标准k-ω SST 湍流模型修正的γ-Re_θ 转捩湍流模型对绕振荡NACA66 水翼流动的数值计算结果与实验结果,对水翼振荡过程的水动力特性和流场结构变化进行了分析研究. 结果表明:与标准k-ω SST 湍流模型的数值计算结果相比,基于标准k-ω SST 湍流模型修正的γ-Re_θ 转捩湍流模型能有效预测绕振荡翼型流场结构和水动力特性,捕捉流场边界层发生的流动分离和转捩现象;绕振荡水翼的流动过程可分为5 个特征阶段,当来流攻角较小时,在水翼前缘发生层流向湍流的转捩现象,水翼动力特征曲线出现变化拐点;随着来流攻角的增大,顺时针尾缘涡逐渐形成并向水翼前缘发展;当攻角较大时,前缘涡分离导致动力失速,水翼的动力特征曲线出现大幅波动;水翼处于顺时针向下旋转阶段,绕水翼的流动状态逐渐由湍流过渡为层流.