Topological analysis of plasma flow control on corner separation in a highly loaded compressor cascade

In this paper, flow behavior and topology structure in a highly loaded compressor cascade with and without plasma aerodynamic actuation (PAA) are investigated. Streamline pattern, total pressure loss coefficient, outlet flow angle and topological analysis are considered to study the effect and mechanism of the plasma flow control on corner separation. Results presented include the boundary layer flow behavior, effects of three types of PAA on separated flows and performance parameters, topology structures and sequences of singular points with and without PAA. Two separation lines, reversed flow and backflow exist on the suction surface. The cross flow on the endwall is an important element for the corner separation. PAA can reduce the underturning and overturning as well as the total pressure loss, leading to an overall increase of flow turning and enhancement of aerodynamic performance. PAA can change the topology structure, sequences of singular points and their corresponding separation lines. Types II and III PAA are much more efficient in controlling corner separation and enhancing aerodynamic performances than type I.     

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.     

机翼尺度效应对等离子体分离流动控制特性的影响

等离子体流动控制技术是一种以等离子体气动激励为控制手段的主动流动控制技术. 为了进一步提高等离子体激励器可控机翼尺度, 以超临界机翼SC(2)-0714大迎角分离流为研究对象, 以对称布局介质阻挡放电等离子体为控制方式, 以测力、粒子图像测速仪为研究手段, 从等离子体激励器特性研究出发, 深入开展了机翼尺度效应对等离子体控制的影响研究, 提出了适用于分离流控制的能效比系数, 探索了分离流等离子体控制机理, 掌握了机翼尺度对分离流控制的影响规律. 结果表明: (1)随着机翼尺度的增大, 布置到机翼上的激励器电极长度会相应增加; 在本文的参数研究范围内, 激励器的平均消耗功率不会随电极长度的增加而线性增大; 当电极长度达到一定阈值时, 激励器的平均消耗功率趋于定值; (2)在固定雷诺数的情况下, 随着机翼尺度的增大, 等离子体的控制效果并未降低, 激励器能效比系数提高; (3)等离子体在主流区诱导的大尺度展向涡与在壁面附近产生的一系列拟序结构成为分离流控制的关键. 研究结果为实现真实飞机的等离子体分离流控制, 推动等离子体流动控制技术工程化应用提供了技术支撑.      

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.      

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.      

低雷诺数翼型蒙皮主动振动气动特性及流场结构数值研究

针对低雷诺数(Re)翼型气动性能差的特点,文章通过对翼型柔性蒙皮施加主动振动的方法,提高翼型低Re下的气动特性,改善其流场结构.采用带预处理技术的Roe方法求解非定常可压缩Navier-Stokes方程,对NACA4415翼型低Re流动展开数值模拟.通过时均化和非定常方法对比柔性蒙皮固定和振动两种状态下的升阻力气动特性和层流分离流动结构.初步研究工作表明在低Re下柔性蒙皮采用合适的振幅和频率,时均化升阻力特性显著提高,分离泡结构由后缘层流分离泡转变为近似的经典长层流分离泡,分离点后移,分离区缩小.在此基础上,文章更加细致研究了柔性蒙皮两种状态下单周期内的层流分离结构及壁面压力系数分布非定常特性和演化规律.蒙皮固定状态下分离区前部流场结构和压力分布基本保持稳定,表现为近似定常分离,仅在后缘位置出现类似于卡门涡街的非定常流动现象.柔性蒙皮振动时从分离点附近开始便产生分离涡,并不断向下游移动、脱落,表现为非定常分离并出现大范围的压力脉动.蒙皮振动使流体更加靠近壁面运动,大尺度的层流分离现象得到有效抑制.      

Separation flow control on a generic ground vehicle using steady microjet arrays

A model of a generic vehicle shape, the Ahmed body with a 25° slant, is equipped with an array of blowing steady microjets 6 mm downstream of the separation line between the roof and the slanted rear window. The goal of the present study is to evaluate the effectiveness of this actuation method in reducing the aerodynamic drag, by reducing or suppressing the 3D closed separation bubble located on the slanted surface. The efficiency of this control approach is quantified with the help of aerodynamic load measurements. The changes in the flow field when control is applied are examined using PIV and wall pressure measurements and skin friction visualisations. By activating the steady microjet array, the drag coefficient was reduced by 9–14% and the lift coefficient up to 42%, depending on the Reynolds number. The strong modification of the flow topology under progressive flow control is particularly studied.     

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.     

Lift enhancement of airfoil and tip flow control for wind turbine

Two techniques that improve the aerodynamic performance of wind turbine airfoils are described. The airfoil S809, designed specially for wind turbine blades, and the airfoil FX60-100, having a higher lift-drag ratio, are selected to verify the flow control techniques. The flow deflector, fixed at the leading edge, is employed to control the boundary layer separation on the airfoil at a high angle of attack. The multi-island genetic algorithm is used to optimize the parameters of the flow deflector. The results indicate that the flow deflector can suppress the flow separation, delay the stall, and enhance the lift. The characteristics of the blade tip vortex, the wake vortex, and the surface pressure distributions of the blades are analyzed. The vortex diffuser, set up at the blade tip, is employed to control the blade tip vortex. The results show that the vortex diffuser can increase the total pressure coefficient of the core of the vortex, decrease the strength of the blade tip vortex, lower the noise, and improve the efficiency of the blade.     

Stall control at high angle of attack with plasma sheet actuators

We analyzed the modifications of the airflow around an NACA 0015 airfoil when the flow was perturbed with electrohydrodynamic forces. The actuation was produced with a plasma sheet device (PSD) consisting in two bare electrodes flush mounted on the surface of the wing profile operated to obtain a discharge contouring the body in the inter-electrode space. We analyze the influence of different parameters of the actuation (frequency, input power, electrodes position) on the aerodynamic performance of the airfoil, basing our study on measurements of the surface pressure distribution and of the flow fields with particle image velocimetry technique. The experiments indicated that at moderate Reynolds numbers (150,000 < Re < 333,000) and at high angles of attack, steady or periodic actuations enabled large improvement of the lift and drag/lift aerodynamic coefficients by reattaching the flow along the extrados. However, to attain the same results steady actuations required larger power consumption. When exciting the flow with a moderate value of non-dimensional power coefficient (ratio of electric power flow with the kinetic power flow), a frequency of excitation produced a peak on the coefficients that evaluate the airfoil performance. This peak in terms of a non-dimensional frequency was close to 0.4 and can be associated to an optimal frequency of excitation. However, our work indicates that this peak is not constant for all stalled flow conditions and should be analyzed considering scale factors that take into account the ratio of the length where the forcing acts and the cord length.     

Mechanism of compressor airfoil boundary layer flow control using nanosecond plasma actuation

The flow control effects of nanosecond plasma actuation on the boundary layer flow of a typical compressor controlled diffusion airfoil are investigated using large eddy simulation method. Three types of plasma actuation are designed to control the boundary layer flow, and two mechanisms of compressor airfoil boundary layer flow control using nanosecond plasma actuation have been found. The plasma actuations located within the laminar boundary layer flow can induce a small vortex structure through influencing on the density and pressure of the flow field. As the small vortex structure moves downstream along the blade surface with the main flow, it can suppress the turbulent flow mixing and reduce the total pressure loss. The flow control effect of the small vortex structure is summarized as wall jet effect. Differently, the plasma actuation located within the turbulent boundary layer flow can act on the shear layer flow and induce a large vortex structure. While moving downstream, this large vortex structure can suppress the turbulent flow mixing too.     

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.     

低Re数下变弯度机翼的非定常气动特性实验研究

利用变弯度机翼模型及相关的风洞实验平台,开展了以弯度变化速率影响为重点的机翼非定常特性研究。实验结果显示,在低Re数(~105)下,机翼弯度非定常变化得到的升阻力系数曲线与准定常条件下的结果存在显著差异。具体表现为:准定常状态下,曲线表现出明显的可逆性;而弯度非定常变化时,曲线在弯度递增区和递减区之间存在明显的迟滞效应,而且随着变形速率的增加,这种迟滞也越明显。流场显示结果表明,这种小St数下出现的流动迟滞是由于弯度变形导致的流动分离的分离点相对机翼运动迟滞所造成的。这说明弯度变化时,分离流场结构的响应时间尺度与弯度变化周期相当,也揭示了该条件下机翼弯度变化对流动的抑制作用主要是通过改变分离区的大小来实现的。     

绕椭球的低速流动研究

理解和预测绕椭球的流动对指导飞行器和潜艇等交通工具的设计具有很强的工程意义. 近年来, 针对椭球绕流开展了大量的实验和数值模拟研究. 对有攻角下椭球绕流分离的定性描述和定量研究, 促进了对三维分离的辨识和拓扑研究. 文章对流场特性进行了分析, 介绍了分离对气动力、噪声、尾迹的影响, 以及实验条件对流动的影响. 上述定常流动与非定常机动过程之间存在明显差异, 非定常机动过程不能作为定常或准定常问题处理, 在机动过程中, 分离出现明显延迟, 气动力出现明显变化. 随后介绍了数值模拟在求解绕椭球流动中的进展, 当前求解雷诺平均的N-S方程湍流模式仍然是解决绕椭球大范围分离流动的主要工程方法, 大涡模拟和分离涡模拟等也逐渐得到了广泛应用. 受限于计算能力, 直接数据模拟只能用于较低雷诺数, 在高雷诺数流动中还不适用. 非定常机动过程的数值模拟较定常状态, 与实验结果的差距要大一些. 最后, 介绍了对椭球绕流场转捩的研究进展, 对T-S转捩与横流转捩的机理和辨识已经较为准确, 数值模拟结果与实验结果基本相符, 但对再附转捩的认识还不够清晰, 尤其是迎风面, 因此椭球绕流转捩的研究还需要依靠实验.      

Active flow control on the 25°Ahmed body using a new unsteady jet

A numerical simulation based on the Large eddy simulation method is carried out on the near wake flow behind a 25° slant angle Ahmed body to analyze and establish a new method to control the near wake flow. An active flow control using a new unsteady jet derived from the traditional synthetic jet is applied to reduce the aerodynamic drag. The control devices are distributed along the separation edges on the rear part of the body. Their effects on the near wake and the rear body by influencing the flow topology and the static pressure distribution are examined respectively. The control frequency of the jet as the key forcing parameter is taken into consideration as well. The different actuation set-ups lead to a max drag reduction of up to 13.6%, which demonstrates a good correlation with the static pressure distribution at the rear end of the body.     

Experimental damping of boundary-layer oscillations using DBD plasma actuators

In the present work artificially excited Tollmien–Schlichting waves were cancelled using plasma actuators operated both in continuous and pulsed modes. To achieve this a vibrating surface, driven by an electromagnetic turbulator, was flush-mounted in a flat plate to excite the TS waves. These were amplified by an adverse pressure gradient induced by an insert on the upper wall of the test section. Control plasma actuators positioned downstream of the excitation actuator attenuate the waves by imparting a steady or unsteady force into the boundary-layer. In the case with steady actuation the two actuators change the velocity profile of the laminar boundary-layer, which then attenuates the waves by itself. In the case of pulsed actuation the actuator creates an unsteady body force to counteract directly the oscillation. As a result the amplitude of the velocity fluctuations at the excitation frequency is reduced significantly in both cases. The principles and the results of the two sets of experiments are presented and discussed.     

Drag reduction on the 25° slant angle Ahmed reference body using pulsed jets

This paper highlights steady and unsteady measurements and flow control results obtained on an Ahmed model with slant angle of 25° in wind tunnel. On this high-drag configuration characterized by a large separation bubble along with energetic streamwise vortices, time-averaged and time-dependent results without control are first presented. The influence of rear-end periodic forcing on the drag coefficient is then investigated using electrically operated magnetic valves in an open-loop control scheme. Four distinct configurations of flow control have been tested: rectangular pulsed jets aligned with the spanwise direction or in winglets configuration on the roof end and rectangular jets or a large open slot at the top of the rear slant. For each configuration, the influence of the forcing parameters (non-dimensional frequency, injected momentum) on the drag coefficient has been studied, along with their impact on the static pressure on both the rear slant and vertical base of the model. Depending on the type and location of pulsed jets actuation, the maximum drag reduction is obtained for increasing injected momentum or well-defined optimal pulsation frequencies.     

Open and closed-loop control of transonic buffet on 3D turbulent wings using fluidic devices

This paper presents an overview of the work performed recently at ONERA on the control of the buffet phenomenon. This aerodynamic instability induces strong wall pressure fluctuations and as such limits aircraft envelope; consequently, it is interesting to try to delay its onset, in order to enlarge aircraft flight envelop, but also to provide more flexibility during the design phase. Several types of flow control have been investigated, either passive (mechanical vortex generators) or active (fluidic VGs, fluidic trailing-edge device (TED)). It is shown than mechanical and fluidic VGs are able to delay buffet onset in the angle-of-attack domain by suppressing the separation downstream of the shock. The effect of the fluidic TED is different, the separation is not suppressed, but the rear wing loading is increased and consequently the buffet onset is not delayed to higher angles of attack, but only to higher lift coefficient. Then, a closed loop control methodology based on a quasi-static approach is defined and several architectures are tested for various parameters such as the input signal, the objective function or, the tuning of the feedback gain. All closed loop methods are implemented on a dSPACE device calculating in real time the fluidic actuators command from the unsteady pressure sensors data.     

圆角化对方柱气动性能影响的流场机理

方形截面柱体的圆角化处理是常用的流动控制方法,但其流场作用机理尚未被澄清.采用大涡模拟方法,在雷诺数为2.2$\times$10$^{4}$时,考虑风攻角的影响,对均匀流作用下的标准方柱和圆角方柱的气动性能和流场特性进行了研究,定量分析了圆角化气动措施和风攻角变化对分离泡特性的影响规律,从流场角度澄清了圆角化气动措施对方柱气动性能的影响机理.研究表明:与标准方柱相比,圆角方柱的表面风压、气动力和涡脱强度呈整体下降的趋势,但圆角方柱的斯特劳哈尔数更高;圆角方柱的"分离泡流态'发生在更小的风攻角范围内,分离泡的出现会进一步造成方柱的尾流变窄,涡脱强度减弱;随着风攻角的增大,分离泡的长度会逐渐减小直至消失,分离泡的中心会逐渐向方柱前角(迎风向)和方柱壁面移动;与标准方柱相比,圆角方柱的气流发生初次分离的位置向下游移动,分离后的剪切层更贴近方柱,因而更易发生再附现象;方柱尾流宽度的减小和涡脱强度的减弱是导致圆角方柱气动力减小和斯特劳哈尔数增大的主要原因.