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.      

Wind tunnel tests of wings at Reynolds numbers below 70 000

 Rectangular planform wings were tested at Reynolds numbers as low as 20 000 in a low turbulence wind tunnel. The lift and drag measurements on a NACA 0012 profile were compared with those for thin flat and cambered plates. For all Reynolds numbers below 70 000 the best profile was a thin plate with a 5% circular arc camber. At all turbulence levels this profile produced the greatest lift-drag ratio, and had the highest lift coefficient at all angles of attack. The 5% camber and all of the thin plates tested were relatively insensitive to either a variation in the Reynolds number, or an increase in the wind tunnel turbulence level, whereas the NACA 0012 was very seriously affected by either, at Reynolds numbers below 50 000. Received: 27 September 1995/Accepted: 21 March 1997     

模型昆虫翼作非定常i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,可产生较大的升力系数.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.     

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.     

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.     

Feedback control of vortex shedding from an inclined flat plate

Open- and closed-loop control of vortex shedding in two-dimensional flow over a flat plate at high angle of attack is numerically investigated at a Reynolds number of 300. Unsteady actuation is modeled as a body force near the leading or trailing edge and is directed either upstream or downstream. For moderate angles of attack, sinusoidal forcing at the natural shedding frequency results in phase locking, with a periodic variation of lift at the same frequency, leading to higher unsteady lift than the natural shedding. However, at sufficiently high angles of attack, a subharmonic of the forcing frequency is also excited and the average lift over the forcing period varies from cycle-to-cycle in a complex manner. It is observed that the periods with the highest averaged lift are associated with particular phase differences between the forcing and the lift, but that this highest-lift shedding cycle is not always stably maintained with open-loop forcing. We design a feedback algorithm to lock the forcing with the phase shift associated with the highest period-averaged lift. It is shown that the compensator results in a stable phase-locked limit cycle for a broader range of forcing frequencies than the open-loop control, and that it is able to stabilize otherwise unstable high-lift limit cycles that cannot be obtained with open-loop control. For example, at an angle of attack of 40°, the feedback controller can increase the averaged magnitude of force on the plate by 76% and increase the averaged lift coefficient from 1.33 to 2.43.     

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.     

Visualization by PIV of dynamic stall on a vertical axis wind turbine

The aerodynamic behavior of a vertical axis wind turbine (VAWT) is analyzed by means of 2D particle image velocimetry (PIV), focusing on the development of dynamic stall at different tip speed ratios. The VAWT has an unsteady aerodynamic behavior due to the variation with the azimuth angle θ of the blade’s sections’ angle of attack, perceived velocity and Reynolds number. The phenomenon of dynamic stall is then an inherent effect of the operation of a VAWT at low tip speed ratios, impacting both loads and power. The present work is driven by the need to understand this phenomenon, by visualizing and quantifying it, and to create a database for model validation. The experimental method uses PIV to visualize the development of the flow over the suction side of the airfoil for two different reference Reynolds numbers and three tip speed ratios in the operational regime of a small urban wind turbine. The field-of-view of the experiment covers the entire rotation of the blade and almost the entire rotor area. The analysis describes the evolution of the flow around the airfoil and in the rotor area, with special focus on the leading edge separation vortex and trailing edge shed vorticity development. The method also allows the quantification of the flow, both the velocity field and the vorticity/circulation (only the results of the vorticity/circulation distribution are presented), in terms of the phase locked average and the random component.     

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.     

MODELING THE UNSTEADY LIFT AND DRAG ON A FINITE-LENGTH CIRCULAR CYLINDER IN CROSS-FLOW

Semi-empirical models for unsteady lift and drag are developed to predict the spectral features of the unsteady forces on a finite-length, right circular cylinder in cross-flow. In general, the models consist of two parts; the spatial variation of r.m.s wall pressure on the cylinder, and the correlation lengths which describe the spatial extent of the correlation of the unsteady wall pressures. Experiments were conducted in a low noise wind tunnel as a function of cylinder diameter Reynolds number (19 200<Re<32 000) and the Strouhal number (0·05< St<3·33), to measure the statistics of the unsteady wall pressures on a model cylinder. These results are incorporated into the theoretical models, and predictions of the spectral characteristics of the lift and drag are made. The r.m.s. wall pressures on the cylindrical surface are found to have the largest amplitude near the cylinder end-cap, and on the rearward portion of the cylinder body. The high levels in these locations are attributed to the separated flow region over the end-cap. The circumferential and axial length-scales decrease exponentially with Strouhal number. Both length-scales exhibit maxima near the Strouhal shedding frequency of St=0·21. The axial length-scales are found to depend on the measurement reference location due to the three-dimensional flow and separated flow region near the end-cap. The unsteady lift and drag predictions using the models developed in this work agree well with previously measured unsteady force data measured on inertial hydrophones exposed to flow. The broadband unsteady lift is found to be greater than the broadband unsteady drag by nominally 3dB.     

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。     

对称翼型低雷诺数小攻角升力系数非线性现象研究

采用Rogers发展的三阶Roe格式,求解非定常不可压N-S方程,时间方向为二阶精度双时间步方法, 数值模拟了对称翼型SD8020低雷诺数(Re=40000,100000)条件下,流场层流分离涡结构和升力系数随攻角的变化．同试验比较证明了数值模拟的正确性．通过对数值模拟时均化流场结果的详细分析,发现对称翼型在小雷诺数0°攻角附近出现的层流分离泡,其内部结构和演化规律都不同于经典层流分离泡模型,从而提出了一种后缘层流分离泡模型．并应用该模型对对称翼型小攻角低雷诺数流场特性以及升力系数非线性效应的形成机理进行了研究和解释．     

Reynolds-number-effects in flow around a rectangular cylinder with aspect ratio 1:5

The paper reports on experiments carried out over a wide range of Reynolds numbers in a high pressure wind tunnel. The model was a sharp-edged rectangular cylinder with aspect ratio height/width 1:5 (width/span ratio 1:10.8), which was investigated in both basic orientations, lengthwise (4×10³<Re<4×10⁵) and perpendicular to the flow (2.7×10⁴<Re<6.4×10⁵). The Reynolds number is based on the height of the model normal to the flow. Steady and unsteady forces were measured with a piezoelectric balance. Thus along with steady (i.e. time averaged values) including the base pressure coefficient, also power spectra and probability density functions were measured yielding for example Strouhal numbers, higher statistical moments, etc. A response diagram for the vortex resonance phenomenon was taken for the natural bending motion of the slender model. If lift coefficient for constant angle of attack is plotted against Reynolds number, a significant Reynolds number effect is seen. For α=4°, the curve shows an inflection point and the lift varies between 0.3 and 0.6. For α=6° and 2° there are similar variations shifted to lower and higher values of Re, respectively. Probably the shapes of separation bubbles that depend on the Reynolds number are responsible for these effects. No Reynolds number effects were observed when the long side was normal to the flow, an orientation where reattachment at the side walls is not possible. Comparing both basic cases (α=0° and 90°), the interpretation of the probability distributions of lift force leads to the conclusion that the possibility of reattachment (α=0°) seems to enhance the degree of order in the vortex shedding process.     

排翼布局飞行器气动性能的实验研究

通过低速低湍流度风洞实验,研究了利用排翼布局改善充气飞机采用大厚度翼型机翼带来的气动效率偏低问题。首先比较了采用不同厚度翼型的单翼与排式双翼布局的气动特性。在此基础上,为了优化排翼布局的气动特性,研究了给后翼安装偏转角对排翼布局气动特性的影响。同时,基于NACA0030翼型,设计了波纹型外形的充气机翼,比较了此外形下单翼和排翼布局气动性能的差异。实验结果表明,采用排翼布局能够改善采用厚翼型单翼布局的气动性能,而给后翼安装一定偏转角可以进一步提高排翼布局的升力和升阻比。采用波纹外形和光滑外形机翼模型的对比结果表明,波纹外形能够在大迎角时改善充气机翼的失速性能。分析认为,造成这一现象的流动机理是由于波纹型机翼在实验条件下提前由层流转捩为湍流,使失速推迟,流动分离现象有所减弱。     

涡轮增压器离心压气机非稳态工况流场分析

利用商用 CFD 软件对一小型车用离心压气机建立了数值模型,并将模拟结果与实验结果进行了对比:稳态的设计转速最高压比相差不超过 0.5%,最高效率相差不超过1.5%;非稳态模拟和实验得到的失速频率均为 3000Hz,模拟结果真实可信.主要利用设计转速下小流量工况时的非稳态数值模拟结果对喘振发生前离心压气机各部件的非稳态流动特点进行了详尽阐述.研究结果表明:小流量工况时离心压气机各部件均出现非稳态流动现象,这种非稳态效应在各部件中表现出不同的特点,且随着流量的减小这种非稳态效应会不断加剧.     

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.     

Dynamic stall model for wind turbine airfoils

A model is presented for aerodynamic lift of wind turbine profiles under dynamic stall. The model combines memory delay effects under attached flow with reduced lift due to flow separation under dynamic stall conditions. The model is based on a backbone curve in the form of the static lift as a function of the angle of attack. The static lift is described by two parameters, the lift at fully attached flow and the degree of attachment. A relationship between these parameters and the static lift is available from a thin plate approximation. Assuming the parameters to be known during static conditions, nonstationary effects are included by three mechanisms: a delay of the lift coefficient of fully attached flow via a second-order filter, a delay of the development of separation represented via a first-order filter, and a lift contribution due to leading edge separation also represented via a first-order filter. The latter is likely to occur during active pitch control of vibrations. It is shown that all included effects can be important when considering wind turbine blades. The proposed model is validated against test data from two load cases, one at fully attached flow conditions and one during dynamic stall conditions. The proposed model is compared with five other dynamic stall models including, among others, the Beddoes–Leishman model and the ONERA model. It is demonstrated that the proposed model performs equally well or even better than more complicated models and that the included nonstationary effects are essential for obtaining satisfactory results. Finally, the influence of camber and thickness distribution on the backbone curve are analysed. It is shown that both of these effects are adequately accounted for via the static input data.     

Experimental analysis of passive bio-inspired covert feathers for stall and post-stall performance enhancement

Technologies inspired by the functioning and behavior of biological beings are commonly developed for aircraft flight. Among the bio-inspired approaches that have grown in interest, particularly for unmanned aerial vehicle flight, is based on the behavior of bird’s cover feathers under higher angles of attack. The covert feathers, when activated by separated flows, promote lift increment that helps in certain maneuvers. This work investigates the benefit in the stall and post-stall performance of employing bio-inspired covert feathers devices attached to an airfoil’s upper surface. To fill the gaps in the recent technical literature, experimental analysis of an SD7003 airfoil was executed in a wind tunnel with the application of bio-inspired covert feathers of different shapes and tapes in three chordwise positions. The bio-inspired devices were conceived to resemble the feathers’ lightness and discrete-distribution along with the wing model. Experiments were carried out measuring the aerodynamic forces and moment at Reynolds number around 170,000 for static and dynamic ramp-up and hold pitching motion. It has been confirmed that the use of bio-inspired covert feathers brought benefits to the stall and post-stall behavior of the airfoil. The maximum lift has increased, and the transition from attached to stalled flow around the airfoil tends to be smoother when the devices were used. Four shapes for the bio-inspired devices and three positions in chordwise direction were considered. The best performance among the case was encountered for a jagged bio-inspired device taped at a quarter-chord position. Indeed, the most forward position for all the devices resulted in higher maximum lift and increment to the respective angle of attack. Ramp-up and hold wind tunnel tests also confirmed the best performance of jagged bio-inspired devices nearer the leading edge. The aerodynamic response to the pitching motion showed that the stall and post-stall regime occur much smoother, indicating that the approach presents good potential for dynamic stall or gust response passive control.

     

Low Reynolds unsteady flow simulation around NACA0012 airfoil with active flow control

This research numerically elucidates the effects of suction and blowing on the enhancement of unsteady aerodynamic characteristics of flows and their corresponding impact on stall delay over the well-known NACA0012 airfoil at various angles of attack (\( 12 \le {\text{AOA}} \le 20 \)) under low Reynolds numbers. For this purpose, an in-house solver written in C++ is developed. The numerical code utilizes the Jameson’s cell-centered finite volume numerical method accompanied by a progressive power-law preconditioning approach to suppress the stiffness of the governing equations. Many numerical simulations are performed over the suction-blowing control parameters, namely, the slot location (\( L_{j} \)), suction/blowing amplitudes (\( A_{j} \)), and suction/blowing angle (\( \theta_{j} \)). Most of the analyses are based on the measurements of the unsteady aerodynamic characteristics behaviors (such as lift, drag, moment coefficients, and stall phenomena) over the airfoil. The numerical results confirm that the unsteady behavior of the flow (vortex shedding) is weakened or approximately removed when suction is used, especially near the leading edge. In all of the test cases, the ratio of the average lift coefficient to the average drag coefficient increases with increasing suction and blowing amplitudes, except in the case of perpendicular blowing. Furthermore, the blowing is more sensitive to the blowing angle compared to the suction. From the suction and blowing results, it is concluded that the former has a more positive impact on the lift and drag characteristics, especially in the case of incompressible flow at Low-Reynolds regimes.     

微型飞行器低雷诺数空气动力学

微型飞行器(MAVs)设计绝不是常规飞行器在尺度上的简单缩小,面临许多技术难题.其中微型飞行器低雷诺数空气动力学是其最为根本的技术瓶颈之一,也是当前受到广泛关注的热点之一.本文紧密结合微型飞行器技术,对这一领域中所面临的低雷诺数空气动力学问题和近两年来该方向国内一些新的进展进行了较为详细的介绍.按照MAVs飞行方式和结构特性进行分类,简单介绍微型飞行器研究中的低$Re$数空气动力学问题.首先介绍了二维和三维固定翼低雷诺数空气动力学问题:包括层流分离泡,翼型升力系数小攻角非线性效应,静态迟滞效应,以及低$Re$数小展弦比机翼气动特性.第2,介绍了拍动翼低雷诺数空气动力学方面的研究工作.包括前人提出的昆虫低$Re$数下获得高升力的多种非定常拍动翼飞行机制:Wagner效应、Weis-Fogh效应(clap-and-fling)、延迟失速效应(delayedstall)、Kramer效应(rotational forces)、尾迹捕获效应(wakecapture)、附加质量效应(addedmass)等.以及国内学者近几年在拍动翼方面取得的一些研究成果.第3,介绍了柔性翼低雷诺数气动问题.研究表明柔性翼对于固定翼微型飞行器提高抗阵风能力,拍动翼微型飞行器产生足够的升力和推力.最后简单介绍了可变形翼(morphingwing)微型飞行器方面的一些研究工作,指出微型飞行器技术可以通过采用可变形翼设计,突破众多的技术瓶颈.另一方面,可变形翼概念可以通过在低成本,低速的MAVs上进行飞行试验,获得非常好的验证平台.