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.      

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.     

On factors influencing arc filament plasma actuator performance in control of high speed jets

Localized arc filament plasma actuators (LAFPAs) have been developed and used at The Gas Dynamics and Turbulence Laboratory for the purpose of controlling high-speed and high Reynolds number jets. The ability of LAFPAs for use in both subsonic and supersonic jets has been explored, and experiments to date have shown that these actuators have significant potential for mixing enhancement and noise control applications. While it has been established that the actuators manipulate instabilities of the jet, the exact nature of how the actuation couples to the flow is still unclear. All of the results previously reported have been based on a nozzle extension that has an azimuthal groove of 1 mm width and 0.5 mm depth along the inner surface approximately 1 mm upstream of nozzle extension exit. The ring groove was initially added to shield the plasma arcs from the high-momentum flow. However, the effect of the ring groove on the actuation mechanism is not known. To explore this effect, a new nozzle extension is designed, which relocates the actuators to the nozzle extension face and eliminates the ring groove. Schlieren images, particle image velocimetry and acoustic results of a Mach 0.9 jet of Reynolds number ~6.1 × 10⁵ show similar trends and magnitudes with and without a ring groove. Thus, it is concluded that the ring groove does not play a primary role in the LAFPAs’ control mechanism. Furthermore, the effect of the duty cycle of the actuator input pulse on the LAFPAs’ control authority is investigated. The results show that the minimum duty cycle that provides complete plasma formation has the largest control over the jet.     

Feedback flow control of a low-Re airfoil by flap actuators

The present paper describes the applicability of the active flow control device, mini electromagnetic flap actuators attached on the leading edge of an airfoil, for the flow separation under both the steady and the unsteady flow conditions in the low Reynolds number region. At first, lift and drag have been measured for a wide variety of the wind speed Reynolds numbers and the angles of attack for the steady flow condition. Then, effects of some simple feedback flow controls, where the time-dependent signal of the lift-drag ratio have been used to detect the stall and served as a trigger to start the actuation, have been explored under the unsteady flow condition for evading the stall. In every low Reynolds number ranging from 30 000 to 80 000, the present actuators worked quite well to delay the stall, increasing in the lift and delaying the stall angle of attack. These aerodynamic modifications by the flap actuators obtained from the steady flow were found to be available even if the manipulation of the actuators started after the stall. Activation threshold of the lift-drag ratio as the input for the feedback control was determined from a stall classification map obtained under the steady flow experiment. Effectiveness of this feedback control was then demonstrated under the condition of the wind speed decrease (Reynolds number from 80 000 to 40 000) keeping the angle of attack constant at 11°, at which the stall occurs without the active control. Immediately after the sudden velocity decrease, the decrease in the lift-drag ratio were detected and the dynamic actuations were successfully started, resulting in evading the stall and keeping high and stable lift. An additional operation of the feedback, in which the running actuation is turned off when the lift-drag ratio shows lower than the second threshold value after operation, was revealed to be effective to keep the high lift force under the condition combined with the wind speed increase and decrease within the low Reynolds number range treated in this study.     

EXPERIMENTAL INVESTIGATION ON LIFT INCREMENT OF A PLASMA CIRCULATION CONTROL AIRFOIL

利用等离子体激励器发展了新型的环量增升技术,并对二维NACA0012翼型绕流实施控制。由于NACA0012翼型为尖后缘构型,环量增升装置由2个非对称型介质阻挡放电等离子体激励器构成。一个等离子体激励器贴附于翼型吸力面靠近后缘处,其诱导的壁面射流沿来流方向指向下游;另一个等离子体激励器贴附于翼型压力面靠近后缘处,其诱导的壁面射流与来流方向相反指向上游。在风洞中通过时间解析二维PIV系统对翼型绕流流场进行了测量,基于翼型弦长的雷诺数Re=20 000。结果表明在等离子体激励器的控制下,翼型压力面靠近后缘处可以形成一个定常回流区,从而起到虚拟气动外形的作用,因此翼型吸力面的流场得到加速,压力面的流场得到减速,使得翼型压力面的吸力以及压力面的压力都得到增加,进而增加了翼型的环量。风洞天平测力实验进一步验证了该环量增升技术的有效性。在整个攻角范围内,施加控制的翼型的升力系数相比没有控制的工况有明显的提高。     

High-fidelity simulations of moving and flexible airfoils at low Reynolds numbers

The present paper highlights results derived from the application of a high-fidelity simulation technique to the analysis of low-Reynolds-number transitional flows over moving and flexible canonical configurations motivated by small natural and man-made flyers. This effort addresses three separate fluid dynamic phenomena relevant to small fliers, including: laminar separation and transition over a stationary airfoil, transition effects on the dynamic stall vortex generated by a plunging airfoil, and the effect of flexibility on the flow structure above a membrane airfoil. The specific cases were also selected to permit comparison with available experimental measurements. First, the process of transition on a stationary SD7003 airfoil section over a range of Reynolds numbers and angles of attack is considered. Prior to stall, the flow exhibits a separated shear layer which rolls up into spanwise vortices. These vortices subsequently undergo spanwise instabilities, and ultimately breakdown into fine-scale turbulent structures as the boundary layer reattaches to the airfoil surface. In a time-averaged sense, the flow displays a closed laminar separation bubble which moves upstream and contracts in size with increasing angle of attack for a fixed Reynolds number. For a fixed angle of attack, as the Reynolds number decreases, the laminar separation bubble grows in vertical extent producing a significant increase in drag. For the lowest Reynolds number considered (Re _c  = 10⁴), transition does not occur over the airfoil at moderate angles of attack prior to stall. Next, the impact of a prescribed high-frequency small-amplitude plunging motion on the transitional flow over the SD7003 airfoil is investigated. The motion-induced high angle of attack results in unsteady separation in the leading edge and in the formation of dynamic-stall-like vortices which convect downstream close to the airfoil. At the lowest value of Reynolds number (Re _c  = 10⁴), transition effects are observed to be minor and the dynamic stall vortex system remains fairly coherent. For Re _c  = 4 × 10⁴, the dynamic-stall vortex system is laminar at is inception, however shortly afterwards, it experiences an abrupt breakdown associated with the onset of spanwise instability effects. The computed phased-averaged structures for both values of Reynolds number are found to be in good agreement with the experimental data. Finally, the effect of structural compliance on the unsteady flow past a membrane airfoil is investigated. The membrane deformation results in mean camber and large fluctuations which improve aerodynamic performance. Larger values of lift and a delay in stall are achieved relative to a rigid airfoil configuration. For Re _c = 4.85 × 10⁴, it is shown that correct prediction of the transitional process is critical to capturing the proper membrane structural response.     

Thrust generation by an airfoil in hover modes

A small airfoil is operated in combined harmonic plunging and pitching motions to generate thrust in a still air environment. By full utilization of dynamic stall vortices large thrust coefficients were attained. The vortical signature of thrust is a simple vortex street with the character of a jet stream.On sabbatical leave from the University of Colorado at Boulder.     

Dynamic stall behavior from unsteady force measurements

A direct force measurement technique employing piezoelectric load cells is used to experimentally investigate a two-dimensional airfoil (NACA 0012) undergoing dynamic stall. The load cells are installed at each end of the airfoil and give the force response in two directions in the plane normal to the airfoil axis during oscillations. Experiments are carried out at a Reynolds number based on the airfoil chord equal to 7.7×10⁴, and at four reduced frequencies, k=0.005, 0.01, 0.02, and 0.04. Phase-averaged lift of the airfoil undergoing dynamic stall is presented. It is observed that hysteresis loops of the lift occur both when the airfoil is pitched to exceed its static stall limit and when it is still within its static stall limit, and they grow in size with increasing k at the same pitching mean angle of attack and pitching amplitude. Both the lift and the drag induced by the pitching motion are further analyzed using the methods of higher order correlation analysis and continuous wavelet transforms to undercover their nonlinear and nonstationary features, in addition to classical FFT-based spectral analysis. The results are quantitatively illustrated by an energy partition analysis. It is found that the unsteady lift and drag show opposite trends when the airfoil undergoes transition from the pre-stall regime to the full-stall regime. The degree of nonlinearity of the lift increases, and the lift show a nonstationary feature in the light-stall regime, while the nonlinearity of the drag decreases, and the drag shows nonstationary feature in both the light-stall and the full-stall regimes. Furthermore, the lift and the drag have significant nonlinear interactions as shown by the correlation analysis in the light-stall regime.     

Development of localized arc filament RF plasma actuators for high-speed and high Reynolds number flow control

Recently developed localized arc filament plasma actuators (LAFPAs) have shown tremendous control authority in high-speed and high Reynolds number flow for mixing enhancement and noise mitigation. Previously, these actuators were powered by a high-voltage pulsed DC plasma generator with low energy coupling efficiency of 5–10%. In the present work, a new custom-designed 8-channel pulsed radio frequency (RF) plasma generator has been developed to power up to 8 plasma actuators operated over a wide range of forcing frequencies (up to 50 kHz) and duty cycles (1–50%), and at high energy coupling efficiency (up to 80–85%). This reduces input electrical power requirements by approximately an order of magnitude, down to 12 W per actuator operating at 10% duty cycle. The new pulsed RF plasma generator is scalable to a system with a large number of channels. Performance of pulsed RF plasma actuators used for flow control was studied in a Mach 0.9 circular jet with a Reynolds number of about 623,000 and compared with that of pulsed DC actuators. Eight actuators were distributed uniformly on the perimeter of a 2.54-cm diameter circular nozzle extension. Both types of actuators coupled approximately the same amount of power to the flow, but with drastically different electrical inputs to the power supplies. Particle image velocimetry measurements showed that jet centerline Mach number decay produced by DC and RF actuators operating at the same forcing frequencies and duty cycles is very similar. At a forcing Strouhal number near 0.3, close to the jet column instability frequency, well-organized periodic structures, with similar patterns and dimensions, were generated in the jets forced by both DC and RF actuators. Far-field acoustic measurements demonstrated similar trends in the overall sound pressure level (OASPL) change produced by both types of actuators, resulting in OASPL reduction up to 1.2–1.5 dB in both cases. We conclude that pulsed RF actuators demonstrate flow control authority similar to pulsed DC actuators, with a significantly reduced power budget.     

Modeling the flow past an oscillating airfoil by the method of viscous vortex domains

The Lagrangian vortex method for solving the Navier-Stokes equations is applied for numerically modeling the unsteady flow past a wing airfoil executing angular oscillations in a viscous incompressible flow. Formulas relating the unsteady forces on the airfoil and the vorticity field are derived. The calculated results are compared with the experimental data for the NACA-0012 airfoil executing harmonic oscillations in an air flow at the Reynolds number Re = 4.4 × 10⁴.     

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.      

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.     

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.     

合成射流对失速状态下翼型大分离流动控制的试验研究

为研究低速状态合成射流在抑制翼型气流分离和推迟失速方面的控制机理, 开展了NACA0021 翼型失速特性射流控制的风洞试验研究. 通过系统性的模型测力、翼型瞬态流场粒子图像测速和边界层速度测定的对比试验, 深入探索了合成射流各参数对翼型失速特性控制效果的影响规律. 试验结果表明射流偏角在翼型升力和失速迎角控制方面的效果对射流动量系数较为敏感: 当动量系数较大时, 近切向射流的控制效果更好. 射流动量系数为0.033 时, 偏角30°的射流使得翼型升力系数峰值提高23.56%, 失速迎角增大5°; 而动量系数较小时, 偏角较大的射流能够获得最佳控制效果. 射流动量系数为0.0026 时, 法向射流对翼型最大升力系数控制效果最好(提高9.2%).      

Validating the URANS shear stress transport γ −Reθ model for low‐Reynolds‐number external aerodynamics

A numerical investigation on low‐Reynolds‐number external aerodynamics was conducted using the transitional unsteady Reynolds‐averaged Navier–Stokes shear stress transport γ ?Re_θ model and the ANSYS‐CFX computational fluid dynamics suite. The NACA 0012 airfoil was exposed to chord‐based Reynolds numbers of 5.0 ×10⁴, 1.0 ×10⁵ and 2.5 ×10⁵ at 0°, 4°and 8°angles of attack. Time‐averaged and instantaneous flow features were extracted and compared with fully turbulent shear stress transport results, XFLR5 panel e ^N method results, and published higher order numerical and experimental studies. The current model was shown to reproduce the complex flow phenomena, including the laminar separation bubble dynamics and aerodynamic performance, with a very good degree of accuracy. The sensitivity of the model to domain size, grid resolution and quality, timestepping scheme, and free‐stream turbulence intensity was also presented. In view of the results obtained, the proposed model is deemed appropriate for modelling low‐Reynolds‐number external aerodynamics and provides a framework for future studies for the better understanding of this complex flow regime. Copyright © 2011 John Wiley & Sons, Ltd.     

The Flow Structure Produced by Pulsed-jet Vortex Generators in a Turbulent Boundary Layer in an Adverse Pressure Gradient

The effect of pulsed jet vortex generators on the structure of an adverse pressure gradient turbulent boundary layer flow was investigated. Two geometrically optimised vortex generator configurations were used, co-rotating and counter-rotating. The duty cycle and pulse frequency were both varied and measurements of the skin friction (using hot films) and flow structure (using stereo PIV) were performed downstream of the actuators. The augmentation of the mean wall shear stress was found to be dependent on the net mass flow injected by the actuators. A quasi steady flow structure was found to develop far downstream of the injection location for the highest pulse frequency tested. The actuator near field flow structure was observed to respond very quickly to variations in the jet exit velocity.     

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.     

Transient phenomena in separation control over a NACA 0015 airfoil

We present the transient phenomena occurring during the impulsive control of flow separation over a NACA0015 airfoil at an incidence angle of 11° and a chord Reynolds number of 1 million. Actuation is performed via pneumatic vortex generators, impulsively activated in order to analyze the transient phenomena corresponding to the attachment process and, conversely, to transient re-separation occurring when the actuators are switched off. Measurements are performed using a linear array of unsteady pressure transducers and a single traversing crosswire. The pressure transducers are positioned in the separated region of the airfoil, which extends ∼ 0.3c upstream of the trailing edge at the above flow condition. To control the flow, the angled fluidic vortex generators are positioned in a single spanwise array located 0.3c downstream of the leading edge of the airfoil. We establish a statistical relationship between pressure and velocity signals during both the uncontrolled steady state and the transient processes of attachment and separation. The unsteady behavior of the attachment process is also qualitatively analyzed via a 0.3 million Reynold number visualizations. The emission of a “starting vortex” is evidenced. This corresponds to a transient increase of drag.     

Numerical investigations into the asymmetric effects on the aerodynamic response of a pitching airfoil

The effects of asymmetric sinusoidal motion on pitching airfoil aerodynamics were studied by numerical simulations for 2-D flow around a NACA0012 airfoil at Re=1.35×10⁵. Various unsteady parameters (amplitude of oscillation, d; reduced frequency, k) were applied to investigate the effect of asymmetry parameter S on the instantaneous force coefficients and flow patterns. The results reveal that S has a noticeable effect on the aerodynamic performance, as it affects the instantaneous force coefficient, maximum lift and drag coefficient, hysteresis loops and the flow structures.     

Large-eddy simulations of plasma-based asymmetric control of supersonic round jets

Localised arc filament plasma actuators are modelled with a validated technique to examine asymmetric control of a perfectly expanded round free jet to deflect its downstream trajectory. The nominal Mach and Reynolds numbers are 1.3 and 1 million, respectively. No-control, symmetrically controlled, and under-expanded jets are also simulated for comparison purposes. Parametric variation of actuation frequency and duty cycle indicate that asymmetric control can alter the trajectory, and, within the confines of the parameters investigated, the optimal forcing scheme was found to correspond to the jet's column-mode frequency and a duty cycle of approximately 60%. Increasing frequency and duty cycle beyond these values have a detrimental effect on control, which is consistent with experimental findings. Asymmetric actuation resulted in significant mixing enhancement on the actuated side, as evidenced by the increased growth rate of the non-dimensional momentum thickness. The effectiveness of control is reduced for under-expanded jet conditions.