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.     

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.     

基于后缘小翼的翼型反流动态失速主动控制试验研究

针对直升机旋翼反流区因反流动态失速导致的非定常载荷、阻力激增以及负升力等问题,开展了基于后缘小翼的翼型反流动态失速主动控制试验研究.采用动态压力测量结合翼型表面压力积分的方法,重点分析了后缘小翼不同的振荡相位差、幅值和减缩频率对反流动态失速控制的影响规律,对比了后缘小翼动态偏转和固定偏转的差异,试验雷诺数Re=3.5×10⁵.结果表明,当后缘小翼与翼型以相同的频率正弦振荡运动,且二者的相位差为0°时,能改善反流动态失速过程中钝几何前缘的流动分离,并在反流状态下实现了翼型负升力系数下降21.2%,阻力系数下降37.5%,俯仰力矩系数迟滞环面积下降44.6%的控制效果;动态偏转的后缘小翼对翼型反流动态失速的控制效果随后缘小翼振荡幅值的增加而增加,但进一步增加振荡幅值对于控制效果的提升有限;当减缩频率增加时,动态偏转的后缘小翼对反流状态下翼型阻力的控制效果会更加明显;后缘小翼的动态偏转与固定偏转都能有效改善翼型在反流中的动态气动性能,但是动态偏转对于不同翼型迎角的适应能力优于固定偏转,并取得了更好的非定常载荷控制以及更好的阻力和负升力改善效果.     

横摆振荡翼型动态气动掠效应试验研究

大型风力机设计对获取翼型更加全面、准确的动态载荷提出更高要求,研究翼型横摆振荡动态气动特性具有重要意义.借助"电子凸轮"技术和动态数据同步采集手段,针对翼型动态"掠效应"首次开展了横摆振荡风洞试验研究,研究表明:横摆振荡翼型的气动曲线存在明显迟滞效应,吸力面压力周期性波动是主要诱因,且随着振荡频率、初始迎角和振幅的增大,气动迟滞特性均增强;升力和压差阻力随横摆角变化的迟滞回线呈"W"形,俯仰力矩迟滞回线呈"M"形,升力差量迟滞回线呈"∞"形;负行程下翼型气动力相对于正行程下的更高,且负行程下翼型气动力随振荡频率的增大而略有增大,正行程下则明显减小;升力系数功率谱密度分布在振荡频率倍频处的能量集中的幅值随着振荡频率增大有增大趋势;吸力面1.2%和40%弦长处压力的滞回特性较强,是由于翼面剪切层涡和动态分离涡周期性发展、运动、破裂和重建;振幅为10?时,升力迟滞曲线呈"∧"形,振幅为30?时,升力迟滞曲线呈"∧∧∧"形.     

弹性振动对翼型气动特性影响的数值模拟

通过求解雷诺平均非定常Navier-Stokes方程,采用数值模拟方法计算了俯仰和沉浮振动对NACA0012翼型平均气动特性的影响.结果表明:对于俯仰运动而言,在迎角13α≤时的升力°和力矩曲线的线性段部分,振幅角的变化对动态平均升力系数和动态平均力矩系数的影响不明显,与静态时的情况基本一致;当迎角14α≥时,翼型振动的平均升力系数和动态平均力矩系数小°于静态时的情况.同一迎角条件下的俯仰振动频率越高时,其动态的平均升力系数和动态平均力矩系数越大,频率较高时的失速迎角相对于频率较低时的情况有所推迟,但相对于静态的失速迎角而言,不同频率下的动态失速迎角均提前.对于沉浮运动而言,动态平均升力系数随振幅和频率的增加而减小,动态失速迎角随振幅和频率的增大而提前.     

等离子体激励气动力学探索与展望

等离子体激励气动力学是研究等离子体激励与流动相互作用下, 绕流物体受力和流动特性以及管道内部流动规律的科学, 属于空气动力学、气体动力学与等离子体动力学交叉前沿领域. 等离子体激励是等离子体在电磁场力作用下运动或气体放电产生的压力、温度、物性变化, 对气流施加的一种可控扰动. 局域、非定常等离子体激励作用下, 气流运动状态会发生显著变化, 进而实现气动性能的提升. 国际上对介质阻挡放电等离子体激励、等离子体合成射流激励及其调控附面层、分离流动、含激波流动等开展了大量研究. 等离子体激励调控气流呈现显著的频率耦合效应, 等离子体冲击流动控制是提升调控效果的重要途径. 发展高效能等离子体激励方法, 通过等离子体激励与气流耦合, 激发和利用气流不稳定性, 揭示耦合机理、提升调控效果, 是等离子体激励气动力学未来的发展方向.      

横摆振荡翼型动态气动掠效应试验研究

大型风力机设计对获取翼型更加全面、准确的动态载荷提出更高要求, 研究翼型横摆振荡动态气动特性具有重要意义. 借助"电子凸轮"技术和动态数据同步采集手段, 针对翼型动态“掠效应”首次开展了横摆振荡风洞试验研究, 研究表明: 横摆振荡翼型的气动曲线存在明显迟滞效应, 吸力面压力周期性波动是主要诱因, 且随着振荡频率、初始迎角和振幅的增大, 气动迟滞特性均增强; 升力和压差阻力随横摆角变化的迟滞回线呈"W"形, 俯仰力矩迟滞回线呈"M"形, 升力差量迟滞回线呈"$\infty$"形; 负行程下翼型气动力相对于正行程下的更高, 且负行程下翼型气动力随振荡频率的增大而略有增大, 正行程下则明显减小; 升力系数功率谱密度分布在振荡频率倍频处的能量集中的幅值随着振荡频率增大有增大趋势; 吸力面1.2%和40%弦长处压力的滞回特性较强, 是由于翼面剪切层涡和动态分离涡周期性发展、运动、破裂和重建; 振幅为$10^{\circ}$时, 升力迟滞曲线呈"$^{\wedge}$"形, 振幅为$30^{\circ}$ 时, 升力迟滞曲线呈"$^{\wedge\wedge\wedge}$"形.      

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。     

等离子体激励器控制圆柱绕流的实验研究

为了发展新型移动附面层控制技术,提升流动控制效率,采用粒子图像测速技术,开展了基于对称布局等离子体气动激励的圆柱绕流控制研究,获得了静止空气下,对称布局激励器诱导流场的演化过程,评估了来流条件下等离子体控制效果,通过等离子体诱导涡实现了虚拟移动附面层控制,分析了诱导涡随时间演化的过程,揭示了圆柱绕流等离子体控制机理.结果表明:(1)在静止空气下,对称布局激励器在刚启动瞬间,会在暴露电极两侧诱导产生一对旋转方向相反的启动涡;随着时间的推移,启动涡逐渐向远离壁面的方向运动;随后,激励器在暴露电极两侧产生了两股速度近似相等,方向相反的诱导射流,诱导射流在柯恩达效应的影响下,朝壁面方向发展.(2)当激励电压峰峰值为19.6 kV,激励频率3kHz时,施加等离子体气动激励后,圆柱脱落涡得到了较好抑制,圆柱阻力系数减小了21.8％;(3)在来流作用下,对称布局激励器在靠近来流一侧,诱导产生了较为稳定的涡结构.诱导涡通过旋转、运动,促进了壁面附近低能气流与主流之间的掺混,抑制了圆柱绕流流场分离,实现了"虚拟移动附面层控制"效果.与传统移动附面层控制技术相比,基于等离子体气动激励的新型移动附面层控制技术不需要复杂、笨重的机构,不会带来额外的阻力,具有潜在的应用前景.      

The onset of dynamic stall revisited

Dynamic stall on a helicopter rotor blade comprises a series of complex aerodynamic phenomena in response to the unsteady change of the blade’s angle of attack. It is accompanied by a lift overshoot and delayed massive flow separation with respect to static stall. The classical hallmark of the dynamic stall phenomenon is the dynamic stall vortex. The flow over an oscillating OA209 airfoil under dynamic stall conditions was investigated by means of unsteady surface pressure measurements and time-resolved particle image velocimetry. The characteristic features of the unsteady flow field were identified and analysed utilising different coherent structure identification methods. An Eulerian and a Lagrangian procedure were adopted to locate the axes of vortices and the edges of Lagrangian coherent structures, respectively; a proper orthogonal decomposition of the velocity field revealed the energetically dominant coherent flow patterns and their temporal evolution. Based on the complementary information obtained by these methods the dynamics and interaction of vortical structures were analysed within a single dynamic stall life cycle leading to a classification of the unsteady flow development into five successive stages: the attached flow stage; the stall development stage; stall onset; the stalled stage; and flow reattachment. The onset of dynamic stall was specified here based on a characteristic mode of the proper orthogonal decomposition of the velocity field. Variations in the flow field topology that accompany the stall onset were verified by the Lagrangian coherent structure analysis. The instantaneous effective unsteadiness was defined as a single representative parameter to describe the influence of the motion parameters. Dynamic stall onset was found to be promoted by increasing unsteadiness. The mechanism that results in the detachment of the dynamic stall vortex from the airfoil was identified as vortex-induced separation caused by strong viscous interactions. Finally, a revised criterion to discern between light and deep dynamic stall was formulated.     

Duty cycle and directional jet effects of a plasma actuator on the flow control around a NACA0015 airfoil

This paper reports on the effects of a series of fluid-dynamic dielectric barrier discharge plasma actuators on a NACA0015 airfoil at high angle of attack. A set of jet actuators able to produce plasma jets with different directions (vectoring effect) and operated at different on/off duty cycle frequencies are used. The experiments are performed in a wind tunnel facility. The vectorized jet and the transient of the flow induced by unsteady duty cycle operation of each actuator are examined and the effectiveness of the actuator to recover stall condition in the range of Reynolds numbers between 1.0 × 10⁵ and 5.0 × 10⁵ (based on airfoil chord), is investigated. The actuator placed on the leading edge of the airfoil presents the most effective stall recovery. No significant effects can be observed for different orientations of the jet. An increase of the stall recovery is detected when the actuator is operated in unsteady operation mode. Moreover, the frequency of the on/off duty cycle that maximizes the stall recovery is found to be a function of the free stream velocity. This frequency seems to scale with the boundary layer thickness at the position of the actuator. A lift coefficient increase at low free stream velocities appears to linearly depend on the supply voltage.     

Exploration of plasma-based control for low-Reynolds number airfoil/gust interaction

Large-eddy simulation (LES) is employed to investigate the use of plasma-based actuation for the control of a vortical gust interacting with a wing section at a low Reynolds number. Flow about the SD7003 airfoil section at 4° angle of attack and a chord-based Reynolds number of 60,000 is considered in the simulation, which typifies micro air vehicle (MAV) applications. Solutions are obtained to the Navier–Stokes equations that were augmented by source terms used to represent body forces imparted by the plasma actuator on the fluid. A simple phenomenological model provided these body forces resulting from the electric field generated by the plasma. The numerical method is based upon a high-fidelity time-implicit scheme and an implicit LES approach which are used to obtain solutions on a locally refined overset mesh system. A Taylor-like vortex model is employed to represent a gust impinging upon the wing surface, which causes a substantial disruption to the undisturbed flow. It is shown that the fundamental impact of the gust on unsteady aerodynamic forces is due to an inviscid process, corresponding to variation in the effective angle of attack, which is not easily overcome. Plasma control is utilised to mitigate adverse effects of the interaction and improve aerodynamic performance. Physical characteristics of the interaction are described, and several aspects of the control strategy are explored. Among these are uniform and non-uniform spanwise variations of the control configuration, co-flow and counter-flow orientations of the directed force, pulsed and continuous operations of the actuator and strength of the plasma field. Results of the control situations are compared with regard to their effect upon aerodynamic forces. It was found that disturbances to the moment coefficient produced by the gust can be greatly reduced, which may be significant for stability and handling of MAV operations.     

Control of the corner separation in a compressor cascade by steady and unsteady plasma aerodynamic actuation

This paper reports experimental results on using steady and unsteady plasma aerodynamic actuation to control the corner separation, which forms over the suction surface and end wall corner of a compressor cascade blade passage. Total pressure recovery coefficient distribution was adopted to evaluate the corner separation. Corner separation causes significant total pressure loss even when the angle of attack is 0°. Both steady and unsteady plasma aerodynamic actuations suppress the corner separation effectively. The control effect obtained by the electrode pair at 25% chord length is as effective as that obtained by all four electrode pairs. Increasing the applied voltage improves the control effect while it augments the power requirement. Increasing the Reynolds number or the angle of attack makes the corner separation more difficult to control. The unsteady actuation is much more effective and requires less power due to the coupling between the unsteady actuation and the separated flow. Duty cycle and excitation frequency are key parameters in unsteady plasma flow control. There are thresholds in both the duty cycle and the excitation frequency, above which the control effect saturates. The maximum relative reduction in total pressure loss coefficient achieved is up to 28% at 70% blade span. The obvious difference between steady and unsteady actuation may be that wall jet governs the flow control effect of steady actuation, while much more vortex induced by unsteady actuation is the reason for better control effect.     

A study on the mechanism of high-lift generation by an airfoil in unsteady motion at low reynolds number

The aerodynamic force and flow structure of NACA 0012 airfoil performing an unsteady motion at low Reynolds number (Re=100) are calculated by solving Navier-Stokes equations. The motion consists of three parts: the first translation, rotation and the second translation in the direction opposite to the first. The rotation and the second translation in this motion are expected to represent the rotation and translation of the wing-section of a hovering insect. The flow structure is used in combination with the theory of vorticity dynamics to explain the generation of unsteady aerodynamic force in the motion. During the rotation, due to the creation of strong vortices in short time, large aerodynamic force is produced and the force is almost normal to the airfoil chord. During the second translation, large lift coefficient can be maintained for certain time period and , the lift coefficient averaged over four chord lengths of travel, is larger than 2 (the corresponding steady-state lift coefficient is only 0.9). The large lift coefficient is due to two effects. The first is the delayed shedding of the stall vortex. The second is that the vortices created during the airfoil rotation and in the near wake left by previous translation form a short “vortex street” in front of the airfoil and the “vortex street” induces a “wind”; against this “wind” the airfoil translates, increasing its relative speed. The above results provide insights to the understanding of the mechanism of high-lift generation by a hovering insect. The project supported by the National Natural Science Foundation of China (19725210)     

High-lift airfoil trailing edge separation control using a single dielectric barrier discharge plasma actuator

Control of flow separation from the deflected flap of a high-lift airfoil up to Reynolds numbers of 240,000 (15 m/s) is explored using a single dielectric barrier discharge (DBD) plasma actuator near the flap shoulder. Results show that the plasma discharge can increase or reduce the size of the time-averaged separated region over the flap depending on the frequency of actuation. High-frequency actuation, referred to here as quasi-steady forcing, slightly delays separation while lengthening and flattening the separated region without drastically increasing the measured lift. The actuator is found to be most effective for increasing lift when operated in an unsteady fashion at the natural oscillation frequency of the trailing edge flow field. Results indicate that the primary control mechanism in this configuration is an enhancement of the natural vortex shedding that promotes further momentum transfer between the freestream and separated region. Based on these results, different modulation waveforms for creating unsteady DBD plasma-induced flows are investigated in an effort to improve control authority. Subsequent measurements show that modulation using duty cycles of 50–70% generates stronger velocity perturbations than sinusoidal modulation in quiescent conditions at the expense of an increased power requirement. Investigation of these modulation waveforms for trailing edge separation control similarly shows that additional increases in lift can be obtained. The dependence of these results on the actuator carrier and modulation frequencies is discussed in detail.     

低雷诺数俯仰振荡翼型等离子体流动控制

针对低雷诺数翼型气动性能差的特点, 通过介质阻挡放电(dielectric barrier discharge, DBD)等离子体激励控制的方法, 提高翼型低雷诺数下的气动特性,改善其流场结构. 采用二维准直接数值模拟方法求解非定常不可压Navier-Stokes方程,对具有俯仰运动的NACA0012翼型的低雷诺数流动展开数值模拟.同时将介质阻挡放电激励对流动的作用以彻体力源项的形式加入Navier-Stokes方程,通过数值模拟探究稳态DBD等离子体激励对俯仰振荡NACA0012翼型气动特性和流场特性的影响.为了进行流动控制, 分别在上下表面的前缘和后缘处安装DBD等离子体激励器,并提出四种激励器的开环控制策略,通过对比研究了这些控制策略在不同雷诺数、不同减缩频率以及激励位置下的控制效果.通过流场结构和动态压强分析了等离子体进行流场控制的机理. 结果表明,前缘DBD控制中控制策略B(负攻角时开启上表面激励器,正攻角时开启下表面激励器)效果最好,后缘DBD控制中控制策略C(逆时针旋转时开启上表面激励器,顺时针旋转时开启下表面激励器)效果最好,前缘DBD控制效果会随着减缩频率的增大而下降, 同时会导致阻力增大.而后缘DBD控制可以减小压差阻力, 优于前缘DBD控制,对于计算的所有减缩频率(5.01~11.82)都有较好的增升减阻效果.在不同雷诺数下, DBD控制的增升效果较为稳定, 而减阻效果随着雷诺数的降低而变差,这是由流体黏性效应增强导致的.      

风力机叶片的三维非定常气动特性估算

结合动量-叶素理论、非定常空气动力和动态失速模型来计算风力机叶片的二维非定常气动特性,并在此基础上经过适当修正后考虑三维旋转效应的非定常气动特性。分析比较二维和三维两种计算结果,给出更为合理的计算叶片非定常气动特性的方法。计算结果表明,风力机叶片的三维非定常气动特性计算结果与二维时的计算结果相比有较大改善。