Delay of finite-span excrescence-induced transition using plasma-based control

Direct numerical simulations are carried out to explore the use of flow control that delays transition generated by excrescence on a plate-like geometry in subsonic flow. Both forward-facing and rearward-facing steps of small roughness heights are considered in the investigation. These are representative of joints and other surface imperfections on wing sections that disrupt laminar flow, thereby increasing skin friction and configuration drag. Unlike previous studies, the steps have a finite lateral extent, such that sharp edges occur in both the spanwise and streamwise directions, and provide a more realistic characterisation of misaligned panels in aerodynamic configurations. The effect of spanwise corners upon transition is examined, and dielectric barrier discharge plasma-based flow control is applied to delay transition and increase the extent of the laminar flow region. Solutions are obtained to the Navier– Stokes equations that were augmented by source terms used to represent body forces imparted by plasma actuators on the fluid. A simple phenomenological model provided these forces resulting from the electric field generated by the plasma. The numerical method is based upon a high-fidelity scheme and an implicit time-marching approach, on an overset mesh system that is used to represent the finite-span steps. Very small-amplitude numerical forcing is employed to generate perturbations, which are amplified by the geometric disturbances and result in transition, similar to the physical situation. Both continuous and pulsed operations of actuators are considered, and the effectiveness of the control is quantified. Transition with the forward-facing step is considerably exacerbated by the presence of a spanwise edge. Plasma control is minimally effective, even with the use of multiple actuators and increased applied force. For the rearward-facing step, transition is substantially delayed by plasma control with small force application.     

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.      

Effect of surface temperature on the stability of the swept attachment line boundary layer

At fairly high Reynolds numbers instability may develop on the line of attachment of the potential flow to the leading edge of a swept wing and lead to a transition to boundary layer turbulence directly at the leading edge [1, 2]. Although the first results relating to the stability and transition of laminar flow at the leading edge of swept wings were obtained almost 30 years ago, the problem remains topical. The stability of the swept attachment line boundary layer was recently investigated numerically with allowance for compressibility effects [3, 4]. Below we examine the effect of surface temperature on the stability characteristics of the laminar viscous heat-conducting gas flow at the leading edge of a side slipping wing.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 6, pp. 78–82, November–December, 1990.     

Reynolds number effects on leading edge vortex development on a waving wing

The waving wing experiment is a fully three-dimensional simplification of the flapping wing motion observed in nature. The spanwise velocity gradient and wing starting and stopping acceleration that exist on an insect-like flapping wing are generated by rotational motion of a finite span wing. The flow development around a waving wing at Reynolds number between 10,000 and 60,000 has been studied using flow visualization and high-speed PIV to capture the unsteady velocity field. Lift and drag forces have been measured over a range of angles of attack, and the lift curve shape was similar in all cases. A transient high-lift peak approximately 1.5 times the quasi-steady value occurred in the first chord length of travel, caused by the formation of a strong attached leading edge vortex. This vortex appears to develop and shed more quickly at lower Reynolds numbers. The circulation of the leading edge vortex has been measured and agrees well with force data.     

Optimization of laminar-turbulent transition delay by means of local heating of the surface

The stability and position of laminar-turbulent transition in the boundary layer on a body heated near the leading edge are analyzed. The point of transition is found using the linear theory of the stability of plane-parallel flow and thee ^N -method. It is shown that by heating a tiny area near the leading edge to a temperature exceeding that of the oncoming flow by a factor of two to four, transition may be delayed, even on a thermally insulated surface. For highly radiating surfaces the energy saved by reducing the friction drag may exceed the heating energy by a factor of three. It is shown that by varying the pressure distribution and surface heating it is possible either to increase the airfoil lift for a fixed transition point or delay transition for a fixed lift.Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 4, pp. 90–99, July–August, 1995.     

基于雨燕翅膀的仿生三角翼气动特性计算研究

针对低雷诺数微型飞行器的气动布局, 设计出类似雨燕翅膀的一组具有不同前缘钝度的中等后掠($\varLambda =50^{\circ}$)仿生三角翼. 为了定量对比研究三角翼后缘收缩产生的气动效应, 设计了一组具有同等后掠的普通三角翼. 为了深入研究仿生三角翼布局的前缘涡演化特性以及总体气动特性, 采用数值模拟方法详细地探索了低雷诺数($Re=1.58\times 10^{4})$流动条件下前缘涡涡流结构和气动力随迎角的变化规律. 分析结果表明, 前缘钝度和后缘收缩对仿生三角翼前缘涡的涡流强度和涡破裂位置有显著影响. 相对于钝前缘来说, 尖前缘使仿生三角翼上下表面的压力差增大, 涡流强度也更大, 增升作用也更显著. 相对于普通三角翼构型, 仿生三角翼的前缘斜切使其阻力更大, 但后缘的收缩使涡破裂位置固定在此位置, 因此整个上翼面保持低压, 总的升力更大. 由于小迎角时升力增大更明显, 因此仿生三角翼的气动效率在小迎角时明显大于普通三角翼. 这些结论对于揭示鸟类的飞行机理以及未来微型仿生飞行器的气动布局设计具有重要的研究价值.      

壁面温度在多参数下对平板边界层的影响研究

为了得到壁面温度在不同来流速度、不同湍流强度条件下对边界层转捩与减阻的影响规律,本文采用Transitionk-kl-ω模型对低来流速度下无压力梯度的光滑平板进行了数值模拟。结果表明,随着来流速度的升高,壁温升高所起到的减阻效果更好,即高来流速度对壁面温度更为敏感。当来流处于中高湍流强度下时,壁温升高能起到推迟转捩的作用,且随着湍流强度的升高,转捩推迟的效果越好,但减阻效果正好相反;当来流处于低湍流强度下时,壁温升高会使得转捩提前发生。壁温升高抑制了边界层内流体的脉动程度,使得层流的稳态不易被破坏,流动更加稳定;同时,壁温升高使得边界层内流体的速度梯度减小,从而降低了壁面摩擦系数,故壁温升高能起到推迟边界层转捩与减阻的作用。     

转捩位置对全动舵面热气动弹性的影响

高超声速附面层的转捩预测一直是流体力学研究中的难点,转捩前后物面的摩擦系数和传热系数会发生改变,转捩位置的不同会影响到飞行器表面热环境,进而使得飞行器的气动弹性特性发生显著变化.鉴于高超声速附面层转捩预测的不确定性,本文分析了转捩位置对高超声速全动舵面热气动弹性的影响.首先分别用层流模型和湍流模型求解N-S方程,得到气动热环境,并对气动热进行参数化;然后在不同转捩位置情况下构造出不同转捩位置的热分布模型,基于此种温度分布,结合热应力和材料属性的影响分析结构的热模态,将结构模态插值到气动网格上,采用基于CFD的当地流活塞理论进行气动弹性分析.以M=6,H=15 km的某舵面为对象进行计算,结果表明:(1)随着转捩位置向后缘移动,结构频率上升,结构颤振速度呈增大趋势,转捩位置的变化能够带来颤振临界速度最大6%的变化量;(2)当转捩位置位于舵轴附近时,结构的颤振特性变化剧烈.通过刚度特性的分解和分析发现,导致颤振特性变化的主要因素在于舵轴的刚度特性变化,舵轴的影响量占整个结构刚度特性变化量的80%以上.     

Effects of a vectored trailing edge jet on delta wing vortex breakdown

Flow visualization was used to study the effects of a vectored trailing edge jet on the leading edge vortex breakdown of a 65° delta wing. The experimental results indicated that there is little effect of the jet on the leading edge vortex breakdown when the angle of the vectored jet is less than 10°. With the increase of the vectored angle ß, the effect of the jet on the flow becomes stronger, i.e., the jet delays the leading edge vortex breakdown in the direction of the vectored jet, and accelerates breakdown of the other leading edge vortex. Moreover, the effect of the jet control tends to be weaker with the angle of attack.     

Correlation of laminar-turbulent transition data over flat plates in supersonic/hypersonic flow including leading edge bluntness effects

The paper discusses flat plate boundary layer transition in supersonic/hypersonic flow conditions. Examination of experimental infrared thermography data illustrates the importance of the leading edge thickness and (non-) uniformity to the transition process. Such observations have triggered the collection of a wide range of experimental data on supersonic/hypersonic flat plate boundary layer transition, and a number of attempts to correlate this data with characteristic parameters including leading edge thickness. Results indicate a strong dependence of the relevant transition parameters on the pressure field in the transition region, as this is determined by the combined effects of leading edge thickness and boundary layer growth/viscous interaction, and particularly on the relative importance of the two effects. In fact, two distinct correlation zones are established, depending on whether the pressure distribution at the onset of transition is dominated by leading edge bluntness effects or by boundary layer growth and viscous interaction, thus limiting the observed data scatter to reasonable levels.Received: 13 August 2002, Accepted: 7 February 2003, Published online: 28 April 2003     

Notes on gas–dynamic design of supersonic flying vehicles

It is shown that the lift–to–drag ratio of a thin delta wing is significantly lower than the lift–to–drag ratio of an infinitely long swept plate with an identical lift force. The effect of sweep on a finite wing may be used by excluding disturbances from the leading edge of the wing via introducing a hardened stream surface (wedge) and increasing the wing length. A three–shock waverider is proposed for choosing the optimal parameters. The sharp wedge may be avoided by replacing planar shock waves by a cylindrical shock wave upstream of the blunted wedge. If the leading edge of the wedge is not parallel to the rib that is a source of the expansion wave, a plate with zero wave drag, generating a lift force, may be obtained behind this rib. The system of regularly intersecting shock waves may be applied to design a forward–swept wing.     

Excitation of controllable perturbations in the three-dimensional boundary layer using plasma actuators

Two versions of the structure of a multi-discharge plasma actuator intended to excite boundary layer perturbations in the neighborhood of the leading swept-wing edge are suggested. The actuator must prevent from appearance and development of the crossflow instability modes leading to laminarturbulent transition under the normal conditions. In the case of flow past a swept wing, excitation of controllable perturbations by the plasma actuator is simulated numerically in the steady-state approximation under the typical conditions of cruising flight of a subsonic aircraft. The local body force and thermal impact on the boundary layer flow which is periodic along the leading wing edge is considered. The calculations are carried out for the physical impact parameters realizable in the near-surface dielectric barrier discharge.     

Wingtip vortex control via the use of a reverse half-delta wing

The effect of a 65° sweep reverse half-delta wing (RHDW), mounted at the squared tip of a rectangular NACA 0012 wing, on the tip vortex was investigated experimentally at Re?=?2.45?×?10⁵. The RHDW was found to produce a weaker tip vortex with a lower vorticity level and, more importantly, a reduced lift-induced drag compared to the baseline wing. In addition to the lift increment, the RHDW also produced a large separated wake flow and subsequently an increased profile drag. The reduction in lift-induced drag, however, outperformed the increase in profile drag and resulted in a virtually unchanged total drag in comparison with the baseline wing. Physical mechanisms responsible for the RHDW-induced appealing aerodynamics and vortex flow modifications were discussed.     

各向异性柔性壁上二维T-S波演化的数值研究

受自然界鸟类羽毛的柔性特征启发, 利用数值模拟的手段进行了各向异性柔性壁面对亚音速边界层中T-S(Tollmien-Schlichting)波空间演化的影响研究. 首先, 刚性壁面上的数值结果与线性理论预测的结果吻合得很好, 验证了所采用的高阶精度格心型有限差分方法的可靠性. 在此基础上, 将部分刚性壁面替换为柔性壁面, 结果表明柔性壁面能够减小甚至消除T-S波的不稳定增长区间, 即抑制T-S波的发展, 因而具有推迟边界层转捩的潜力. 柔性壁面的变形不仅有对应T-S波波形的成分, 还会因柔性段前缘引起波长更长, 与T-S波频率相同的壁面波动. 随后开展的参数研究表明, 增大壁面阻尼削弱了前缘引起的壁面波动; 增大壁面的刚度、张力以及弹性系数都会使得壁面的刚性增强, 整体变形幅度下降; 柔性壁面的支撑杠杆臂倾角越大, 壁面刚性越强. 以上参数的增大均会使得柔性壁面抑制T-S波的效果降低. 此外, 当流动反方向流过时, 抑制T-S波的效果也会明显下降. 这些研究结果旨在揭示鸟类高效飞行的部分奥秘, 为被动减阻提供新的思路.      

基于MEMS的流动主动控制技术及其研究进展

MEMS技术与流动控制技术的结合, 使得流动主动控制技术的实际应用逐步成为现实, 极大地推动了流动主动控制技术的发展. 简述了流动主动控制技术的基本原理、关键技术, 以及应用MEMS技术实现流动主动控制的机理和途径. 介绍了几家国外研究机构近年来在流动主动控制技术领域基于MEMS技术的微传感、微控制和微执行技术及其集成技术的研究进展, 以及在三角翼前沿涡控制、减阻研究、发动机喷流控制、细长体背风面分离流控制等方面的应用情况.      

Active compliant wall for skin friction reduction

In order to reduce skin friction drag, an active laminarisation method is developed. Laminar-turbulent boundary layer transition caused by Tollmien–Schlichting (TS) waves is delayed by attenuation of these convective instabilities. An actively driven compliant wall is integrated as part of a wing’s surface. Different configurations of piezo-based actuators are combined with an array of sensitive surface flow sensors. Wall-normal actuation as well as inclined wall displacement are investigated. Together with a realtime-control strategy, transition onset is shifted downstream by six average TS-wave lengths. Using the example of flow velocity, the influence of variable flow conditions on TS-damping rates was investigated. Besides, the boundary layer flow downstream of the active wall area as well as required wall deflections and the global damping effect on skin friction are presented in this paper.     

上下反翼对双后掠乘波体低速特性的影响

相比于传统乘波体外形, 双后掠乘波体在保持高超声速良好性能的条件下能够提升乘波体低速气动性能, 但其仍存在低速稳定性不好等缺陷. 本文从密切锥乘波体理论提出给定前缘型线的乘波体设计方法, 通过给定三维前缘型线分别生成具有相同平面投影形状的上反和下反机翼双后掠乘波体. 使用CFD技术评估不同上下反程度外翼乘波体的低速性能, 分析升阻特性以及流场涡结构特点. 选取稳定性判据, 研究上下反翼对纵向和横侧向稳定性的影响. 结果表明, 机翼上下反对乘波体低速升阻特性影响较小; 不同外形均为纵向静不稳定的, 且俯仰力矩变化趋势比较类似, 机翼下反可使气动焦点位置后移, 提升纵向稳定性; 机翼上反有助于提升乘波体的横向静稳定性, 而下反则会下降; 机翼上反可以提升侧向稳定性, 且上反程度越大提升效果越明显; 同时机翼上反使乘波体的偏航动态稳定性有明显提升, 下反则会降低, 影响程度与机翼上下反程度呈正相关. 通过结果分析, 说明通过机翼上下反改善乘波体低速稳定性是可行的, 为乘波体在宽速域高超声速飞行器中的应用拓展了途径.      

HEAT-PIPE-COOLED THERMAL PROTECTION FOR THE LEADING EDGE OF THE WING

针对高超声速飞行器飞行时翼前缘存在着严重的气动加热问题，为了保证翼前缘的尖锐外形，提出疏导式热防护结构，利用内置高温热管结构为翼前缘提供热防护。采用数值模拟和电弧风洞试验的方法对翼前缘疏导式结构进行了分析，得到翼前缘内置高温热管具有的防热效果。数值模拟结果表明在一定热环境条件下，翼前缘驻点温度下降了304K，尾部最低温度升高了130K，实现了热流从高温区到低温区的疏导，减弱了翼前缘的热载荷，强化了翼前缘的热防护能力。通过电弧风洞试验可以获得相同的热防护结果，并且在一定飞行条件下高温热管可以自适应启动，验证了数值模拟方法的准确性以及翼前缘内置高温热管疏导式热防护结构的可行性。     

Wings of minimum drag for given lift

The variational problem of determining the optimal shape (camber and twist) of the midsurface of a wing having minimum wavedrag is examined in the linear formulation. It is shown that for wings with supersonic leading edge and straight trailing edge, whose shape is given in the form of a double polynomial, the over-all aerodynamic characteristics can be simply expressed in terms of the equation for the leading edge of the wing. This makes it possible not only to solve the variational problem by the Ritz method and obtain the minimum wave drag [1] but also to find the optimal shape of the wing. As examples we consider delta and double-delta wings.     

Receptivity of the inclined-cylinder attachment-line boundary layer to vortex perturbations

The receptivity of the boundary layer in the neighborhood of the attachment line of a cylinder inclined to the flow with respect to periodic vortex perturbations frozen into the stream is investigated. The problem considered simulates the interaction between external turbulence and the leading-edge swept wing boundary layer. It is shown that if the direction of the external perturbation vector is almost parallel to the leading edge, then the external perturbations are considerably strengthened at the outer boundary layer edge. This effect can cause laminar-turbulent transition on the attachment line at subcritical Reynolds numbers.Translated from Izvestiya Rossiiskoi Academii Nauk, Mekhanika Zhidkosti i Gaza, No. 6, 2004, pp. 72–85. Original Russian Text Copyright © 2004 by Ustinov.