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.      

Transonic flow past an airfoil with mini-flaps

The effect of mini-flaps located on either the lower or upper side of an airfoil near its trailing edge on the flow around the trailing edge and the global flow past the airfoil is numerically investigated. The flow pattern near the trailing edge is compared with that on which the Chaplygin-Joukowski hypothesis is based. The mini-flap effect on the aerodynamic characteristics of the airfoil is studied.     

钝后缘风力机翼型的环量控制研究

钝后缘风力机翼型具有结构强度高、对表面污染不敏感等优点,但其较大的阻力系数使得翼型的整体气动特性不够理想. 利用环量控制方法对钝后缘风力机翼型进行了流动控制,以改善钝后缘风力机翼型的气动特性,减弱尾迹区脱体涡强度. 通过对钝后缘风力机翼型环量控制方法进行相关的数值模拟,对比研究了环量控制方法的增升减阻效果, 研究了环量控制下翼型升阻力特性随射流动量系数的变化规律,并对不同射流动量系数下环量控制方法的气动品质因子和控制效率进行了分析. 研究结果表明:环量控制方法能够大幅提升钝后缘风力机翼型的升力系数,同时有效地降低翼型的阻力系数; 翼型的升力系数随射流动量系数的增大而增大,表现出很明显的分离控制阶段和超环量控制阶段的变化规律; 射流能耗的功率系数随射流动量系数的增大而增大,且增长速率逐渐增大;实施环量控制方法后叶片的输出功率同样随射流动量系数增大而增大,但增长速率逐渐降低. 总体来说,环量控制方法可以有效地改善钝后缘风力机翼型的气动特性以及功率输出特性,在大型风力机流动控制中具有很好的应用前景.      

Measurement of tonal-noise characteristics and periodic flow structure around NACA0018 airfoil

The characteristics of tonal noise and the variations of flow structure around NACA0018 airfoil in a uniform flow are studied by means of simultaneous measurement of noise and velocity field by particle-image velocimetry to understand the generation mechanism of tonal noise. Measurements are made on the noise characteristics, the phase-averaged velocity field with respect to the noise signal, and the cross-correlation contour of velocity fluctuations and noise signal. These experimental results indicate that the tonal noise is generated from the periodic vortex structure on the pressure surface of the airfoil near the trailing edge of the airfoil. It is found that the vortex structure is highly correlated with the noise signal, which indicates the presence of noise-source distribution on the pressure surface. The vorticity distribution on the pressure surface breaks down near the trailing edge of the airfoil and forms a staggered vortex street in the wake of the airfoil.     

Numerical investigation on effects of induced jet on boundary layer and turbulent models around airfoils

In this study the effects of induced jet at trailing edge of a two dimensional airfoil on its boundary layer shape, separation over surface and turbulent parameters behind trailing edge are numerically investigated and compared against a previous experimental data. After proving independency of results from mesh size and obtaining the required mesh size, different turbulent models are examined and RNG k-epsilon model is chosen because of good agreement with experimental data in velocity and turbulent intensity variations. A comparison between ordinary and jet induced cases, regarding numerical data, is made. The results showed that because of low number of measurement points in experimental study, turbulent intensity extremes are not captured. While in numerical study, these values and their positions are well calculated and exact variation of turbulent intensity is acquired. Also a study in effect of jet at high angles of attack is done and the results showed the ability of jet in controlling separation and reducing wake region.     

Investigation of the control of the flow around a wing using mini-flaps

The flow around an airfoil with a mini-flap mounted on the lower or upper wing surface is investigated. The results are obtained by measuring the pressure distribution over the airfoil surface and the forces acting on it for Mach and Reynolds numbers M = 0.1?0.8 and Re = (0.6?3.8) × 10⁶. It is shown that, as distinct from known devices such as Gurney flaps, blunt trailing edge, etc., for controlling the flow in the vicinity of the trailing edge of an airfoil, a mini-flap mounted on the undersurface produces gas flow from the upper to the lower surface around a sharp edge. In this case the flow pattern is considerably affected not only near the trailing edge but also over the entire airfoil. The pressure redistribution over the airfoil makes it possible to increase or decrease the wing lift. Thanks to the low hinge moment, the mini-flap can serve as an effective means of low-inertia control of the flow around a wing.     

Mathematical Modeling of Time-Dependent Aerodynamic Loads on an Airfoil with Dynamic Deflection of a Control on the Trailing Edge

The linear problem of the time-dependent inviscid flow past a thin symmetric airfoil with a control on its trailing edge deflected in accordance with an arbitrary law is considered. The aerodynamic loads on the airfoil are calculated. The intensity of the vortex wake shed from the airfoil is determined by numerically solving a Volterra integral equation of the first kind. Questions of the mathematical modeling of the time-dependent aerodynamic loads in a form convenient for the joint solution of the problems of aerodynamics and flight dynamics are also considered. The results of the modeling are compared with the numerical solutions obtained.__________Translated from Izvestiya Rossiiskoi Academii Nauk, Mekhanika Zhidkosti i Gaza, No. 3, 2005, pp. 157–169.Original Russian Text Copyright © 2005 by Khrabrov.     

On bimodal flutter behavior of a flexible airfoil

The dynamic aeroelastic behavior of an elastically supported airfoil is studied in order to investigate the possibilities of increasing critical flutter speed by exploiting its chord-wise flexibility. The flexible airfoil concept is implemented using a rigid airfoil-shaped leading edge, and a flexible thin laminated composite plate conformally attached to its trailing edge. The flutter behavior is studied in terms of the number of laminate plies used in the composite plate for a given aeroelastic system configuration. The flutter behavior is predicted by using an eigenfunction expansion approach which is also used to design a laminated plate in order to attain superior flutter characteristics. Such an airfoil is characterized by two types of flutter responses, the classical airfoil flutter and the plate flutter. Analysis shows that a significant increase in the critical flutter speed can be achieved with high plunge and low pitch stiffness in the region where the aeroelastic system exhibits a bimodal flutter behavior, e.g., where the airfoil flutter and the plate flutter occur simultaneously. The predicted flutter behavior of a flexible airfoil is experimentally verified by conducting a series of systematic aeroelastic system configurations wind tunnel flutter campaigns. The experimental investigations provide, for each type of flutter, a measured flutter response, including the one with indicated bimodal behavior.     

An experimental investigation on the surface water transport process over an airfoil by using a digital image projection technique

In the present study, an experimental investigation was conducted to characterize the transient behavior of the surface water film and rivulet flows driven by boundary layer airflows over a NACA0012 airfoil in order to elucidate underlying physics of the important micro-physical processes pertinent to aircraft icing phenomena. A digital image projection (DIP) technique was developed to quantitatively measure the film thickness distribution of the surface water film/rivulet flows over the airfoil at different test conditions. The time-resolved DIP measurements reveal that micro-sized water droplets carried by the oncoming airflow impinged onto the airfoil surface, mainly in the region near the airfoil leading edge. After impingement, the water droplets formed thin water film that runs back over the airfoil surface, driven by the boundary layer airflow. As the water film advanced downstream, the contact line was found to bugle locally and developed into isolated water rivulets further downstream. The front lobes of the rivulets quickly advanced along the airfoil and then shed from the airfoil trailing edge, resulting in isolated water transport channels over the airfoil surface. The water channels were responsible for transporting the water mass impinging at the airfoil leading edge. Additionally, the transition location of the surface water transport process from film flows to rivulet flows was found to occur further upstream with increasing velocity of the oncoming airflow. The thickness of the water film/rivulet flows was found to increase monotonically with the increasing distance away from the airfoil leading edge. The runback velocity of the water rivulets was found to increase rapidly with the increasing airflow velocity, while the rivulet width and the gap between the neighboring rivulets decreased as the airflow velocity increased.     

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.     

A new direct design method for the medium thickness wind turbine airfoil

The newly developed integral function of airfoil profiles based on Trajkovski conformal transform theory could be used to optimize the profiles for the thin thickness airfoil. However, it is hard to adjust the coefficients of the integral function for the medium thickness airfoil. B-spline curve has an advantage of local adjustment, which makes it to effectively control the airfoil profiles at the trailing edge. Therefore, a new direct design method for the medium thickness wind turbine airfoil based on airfoil integral expression and B-spline curve is presented in this paper. An optimal mathematical model of an airfoil is built. Two new airfoils with similar thickness, based on the new designed method and the original integral method, are designed. According to the comparative analysis, the CQU-A25 airfoil designed based on the new method exhibits better results than that of the CQU-I25 airfoil which is designed based on the original method. It is demonstrated that the new method is feasible to design wind turbine airfoils. Meanwhile, the comparison of the aerodynamic performance for the CQU-A25 airfoil and for the DU91-W2-250 airfoil is studied. Results show that the maximum lift coefficient and the maximum lift/drag ratio of the CQU-A25 airfoil are higher than the ones of DU91-W2-250 airfoil in the same condition. This new airfoil design method would make it possible to design other airfoils with different thicknesses.     

Energy harvesting of a heaving and forward pitching wing with a passively actuated trailing edge

This study experimentally investigates the energy harvesting capabilities of an oscillating wing with a passively actuated trailing edge. The oscillation kinematics are composed of a combined heaving and forward pitching motions, where the pitching axis is well behind the wing center of mass. Passive actuation is attained by connecting the trailing edge with the wing body using a torsion rod. The degree of flexibility of the trailing edge is represented by the Strouhal number based on the trailing edge natural frequency. The trailing edge passive response is studied for oscillation Strouhal numbers of 0.017, 0.025 and 0.033. Instantaneous aerodynamic forces are measured in a closed loop wind tunnel at a Reynolds number of 40 000, based on the free stream velocity and the wing chord length. Measured results include the effective angle of attack induced by the trailing edge actuation as well as the lift and moment during the oscillation cycle. For the imposed kinematics in this study, the pitching motion has a positive contribution to the mean power output whereas the heaving motion has a relatively small but negative contribution. Additionally, by decreasing the natural frequency of the trailing edge closer to that of the imposed oscillation frequency, the magnitude of the lift and moment forces and hence the mean power output, increases. It is found that there exists a strong correlation between mean power output and the effective angle of attack, shown through the passive trailing edge response, resulting in an increase in energy harvesting potential.     

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。     

A Model Problem of the Design and Aerodynamical Calculation of the Gliding Wing Airfoil of a Wing-in-Ground Vehicle

The problem of the design and aerodynamical calculation of a wing airfoil whose trailing edge glides over a plane horizontal surface is formulated and solved. The known airfoil undersurface is a rectilinear segment forming a given angle with the ground surface, while the upper surface is sought on the basis of a preassigned velocity distribution. This distribution is taken from a class of hydrodynamically advantageous distributions ensuring separationless flow around the airfoil within the framework of the mathematical flow model adopted. The problem is reduced to a mixed boundary value problem in a half-plane and solved analytically. For calculating the lift coefficient the assumption of a thin jetlet flowing between the horizontal region of the airfoil contour and the ground surface is introduced. The effect of the law of pressure decrease along the jetlet, from the stagnation to the exit value, on the lift coefficient is studied. On the basis of the calculations performed a conclusion is drawn concerning the influence of the rectilinear region slope on the airfoil shape; it is also shown how the slope and the value of the maximum velocity on the airfoil affect its shape and the lift coefficient.     

Maneuvering Aspects and 3D Effects of Active Airfoil Flow Control

An active flow control experiment was conducted on a cropped NACA 0018 airfoil to study 3D effects and maneuverability aspects made possible by a segmented actuation system installed in the airfoil. The 14 piezo-fluidic actuators were installed at the corner of the cropped region, inclined at 30° to the local surface, facing downstream. Operating all actuators at unison significantly increased lift and generated a pitch-down moment. Operating all actuators at the same magnitude but varying the phase along the span generated larger lift-increment, with respect to the uniform phase excitation. Significant rolling moment can be generated when only half-span of the wing is actuated. The latter effect, as indicated by the 3D pressure distribution, persists to the leading edge even though the excitation was introduced close to the trailing edge. When a pair, out of the possible fourteen actuators is not operating, very little of the control authority is lost. This is an important finding when issues like fault tolerance and robustness of fluidic-piezo actuators are considered.     

Experimental study of pitching and plunging airfoils at low Reynolds numbers

Measurements of the unsteady flow structure and force time history of pitching and plunging SD7003 and flat plate airfoils at low Reynolds numbers are presented. The airfoils were pitched and plunged in the effective angle of attack range of 2.4°–13.6° (shallow-stall kinematics) and ?6° to 22° (deep-stall kinematics). The shallow-stall kinematics results for the SD7003 airfoil show attached flow and laminar-to-turbulent transition at low effective angle of attack during the down stroke motion, while the flat plate model exhibits leading edge separation. Strong Re-number effects were found for the SD7003 airfoil which produced approximately 25 % increase in the peak lift coefficient at Re = 10,000 compared to higher Re flows. The flat plate airfoil showed reduced Re effects due to leading edge separation at the sharper leading edge, and the measured peak lift coefficient was higher than that predicted by unsteady potential flow theory. The deep-stall kinematics resulted in leading edge separation that led to formation of a large leading edge vortex (LEV) and a small trailing edge vortex (TEV) for both airfoils. The measured peak lift coefficient was significantly higher (~50 %) than that for the shallow-stall kinematics. The effect of airfoil shape on lift force was greater than the Re effect. Turbulence statistics were measured as a function of phase using ensemble averages. The results show anisotropic turbulence for the LEV and isotropic turbulence for the TEV. Comparison of unsteady potential flow theory with the experimental data showed better agreement by using the quasi-steady approximation, or setting C(k) = 1 in Theodorsen theory, for leading edge–separated flows.     

On the Vortex Dynamic of Airflow Reattachment Forced by a Single Non-thermal Plasma Discharge Actuator

Commercial and military aircrafts or miniature aerial vehicles can suffer from massive flow separation when high angles of attack are required. Single dielectric barrier discharge (DBD) actuators have demonstrated their capability of controlling such a separated flow at low external velocity. However, the processes resulting in the improvement of the flight performances remain unclear. In the present study, the reattachment process along the suction side of a NACA 0015 placed at an angle of attack of 16° is experimentally investigated for an external velocity of 20 m/s (Re = 260,000). A single DBD actuator is mounted at the leading edge of the model. The velocity fields above the suction side of the airfoil are measured by a high-speed acquisition system (3 kHz). The results indicate that the baseline flow presents shed vortices that form at the leading edge and linearly grow along the free shear layer axis. This vortex shedding is organized and exhibits a specific frequency of about 90 Hz. The continuous actuation produces a partial flow reattachment up to 70% of the chord length. Temporal cross-correlation function indicates the presence of a vortex shedding at the trailing edge of the controlled flow. Finally, the temporal analysis demonstrates that the reattachment process requires 50 ms to reach a stabilized attached flow. The time-resolved analysis of the reattachment suggests that the actuation by plasma discharge acts as a catalyser by reinforcing one of the coherent flow structures already existing in the natural flow.     

Numerical solution of the problem of the mixing of the boundary layers shed from the trailing edge of a wing

During the mixing of viscous incompressible flows with different velocities, in the vicinity of a trailing edge an interaction region with a three-layer structure is formed, similar to that in the case of symmetric shedding with equal velocities. The boundary layers developing on the upper and lower sides of the airfoil form a viscous mixing layer, or vortex sheet, which separates the flows downstream of the trailing edge. The boundary value problem corresponding to the flow in the viscous sublayer in the vicinity of the trailing edge of a flat plate is solved for high Reynolds numbers using an efficient numerical method for solving the equations of asymptotic interaction theory.     

The physics of modeling unsteady flaps with gaps

Research indicates that active control concepts have promise in mitigating numerous adverse phenomena associated with the aeromechanics of lifting surfaces. These techniques are being applied to delay stall of fixed wing aircraft, as well as to eliminate or mitigate vibratory loads, blade–vortex interaction, and dynamic stall of the flow about rotorcraft and wind turbine blades. These phenomena are nonlinear and unsteady for dynamic systems, which add yet another layer of complexity on the physics of the flow. While a plethora of different active control techniques is being explored, the use of trailing edge flaps appears to be one of the more viable and cost-effective concepts. Static multi-element airfoils and wings have been analyzed computationally, but little exists on the ability to model these when the airfoil and flap are dynamic. The costs associated with modeling the gap between the airfoil and flap have led to approximations where the flap is modeled only as a morphed tip of the airfoil (no gap). Using a hybrid Reynolds-Averaged Navier–Stokes/Large-Eddy-Simulation turbulence technique, an oscillating flapped airfoil has been studied to determine the influence of modeling the gap on the performance and acoustic signature of the airfoil. Results are compared with the experimental data to confirm the validity of the computational approach. Both attached and separated (dynamic stall) oscillating flows are examined. The physics within the gap are found to be important for the airfoil performance when stall is encountered, as well as when acoustic signatures are required.     

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。