Receptivity of the Swept Wing Boundary Layer to a Steady Flow Inhomogeneity

The boundary layer on a plate with an inclined blunt leading edge is investigated for a free-stream flow with a small span-periodic velocity inhomogeneity. This flow simulates the penetration of the outer turbulence into the swept wing boundary layer. It is shown that the boundary layer perturbations generated by the inhomogeneity generally have a streamwise velocity component significantly greater than the initial inhomogeneity amplitude. The dependence of the perturbations on the distance from the leading edge and the spanwise inhomogeneity period is found. It is shown that the swept wing boundary layer is less sensitive to the perturbation type in question than the straight wing boundary layer.     

Experimental transition delay using glow-discharge plasma actuators

In the present work plasma actuators were applied in a flat-plate boundary layer with an adverse pressure gradient to influence the transition of the boundary layer. The first actuator downstream of the leading edge is operated in pulsed mode to introduce perturbations into the boundary layer to promote transition. Two steady operating actuators further downstream damp the perturbations significantly, which results in transition delay.     

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.     

Stability and Receptivity of the Swept-Wing Boundary-Layer Flow Controlled by a Periodic Sequence of Plasma Actuators

Fluid Dynamics - The process of flow control in the boundary layer on a swept wing using a span-periodic sequence of plasma actuators mounted at an angle to the leading edge is modeled. The...     

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.     

Linear evolution of controlled disturbances in the supersonic boundary layer on a swept wing

The linear development of controlled disturbances in the three-dimensional supersonic boundary layer on a swept model wing with a sharp leading edge is experimentally investigated at the Mach number 2. The spatial-temporal and spectral-wave characteristics of the wave train of unstable disturbances are obtained. The asymmetry of these characteristics, due to the secondary flow in the three-dimensional boundary layer, is confirmed.     

Acoustic Field Effect on Laminar Turbulent Transition on a Swept Wing in the Favorable Pressure Gradient Region

The development of disturbances in a three-dimensional boundary layer on a swept wing model is studied both under natural conditions and for artificial excitation of traveling waves by an acoustic field. It is found that steady-state streamwise structures are formed in the three-dimensional boundary layer; under natural conditions a wave packet leading to turbulence is detected. When the flow is exposed to the action of an acoustic field at a frequency from the wave packet, disturbances whose velocity along the streamwise structures is equal to 0.55 of the oncoming flow velocity are formed, while the laminar-turbulent transition is displaced upstream.     

Effect of compressibility on plasma-based transition control for a wing with leading-edge excrescence

Compressibility effects were numerically investigated for use of plasma-based flow control, which was applied to delay transition generated by excrescence on the leading edge of a wing. The wing airfoil section incorporates a geometry that is representative of modern reconnaissance air vehicles, and has an appreciable region of laminar flow at design conditions. Modification of the leading edge can be caused by the accumulation of debris, insect impacts, microscopic ice crystal formation, damage, or structural fatigue, resulting in premature transition and an increase in drag. A dielectric barrier discharge (DBD) plasma actuator, located downstream of the excrescence, was employed to delay transition, mitigate the effects of turbulence, decrease drag, and increase energy efficiency. Solutions were obtained for several Mach numbers, up to the transonic range. The effect of compressibility on transitional behaviour was explored, and the effectiveness of plasma-based control to delay transition with increasing Mach number was determined.     

Robustness of the Leading Edge Vortex on Rotating Wings to Unsteady Perturbations

The vast majority of research works on low aspect ratio rotating wings report that, at high angle of attack, the leading edge vortex that forms on the upper surface of the wing is stable. This ‘trick’ is used by insects and auto-rotating seeds, for example, to achieve the desirable amount of lift. Yet, a few experimental studies suggest that leading edge vortices might be unstable under similar, low Rossby number, conditions. While it is unclear what causes vortex shedding in these studies, the present communication explores the sensitivity of leading edge vortex attachment to perturbations of the rotating speed and demonstrates that shedding can be triggered even for very small perturbations, corresponding to wing tip displacements lower than 1% of the wing chord.

     

Propagation of perturbations upstream with interaction between a hypersonic flow and a boundary layer

It is shown that, for hypersonic flows with moderate and strong degrees of interaction, perturbations brought about, for example, by a bottom opening or by any other sort of obstacle are propagated up to the leading edge of a solid body. Local regions with very large pressure gradients cannot arise in the flow. This is connected with the possibility of the development of breakaway zones with a length on the order of magnitude of the size of the solid body, described in the first approximation by the equations of the boundary layer. From a mathematical point of view the problem comes down to establishing the nonsingular nature of the solution near the leading edge, and to finding eigensolutions which make it possible to satisfy the boundary conditions at the trailing edge of the solid body. It is shown that, with a weak interaction between the hypersonic flow and the boundary layer, there may arise short flow regions with free interaction and locally nonviscous flows with large pressure gradients, within the limits of which the perturbations may move upstream.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 4, pp. 40–49, July–August, 1970.In conclusion, the author thanks V. V. Sychev for his evaluation of the problem.     

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。     

Investigation of the Boundary Layer Behavior In the Vicinity of the Attachment Line on a Yawed Wing and the Possibilities of Deturbulization

The results of an experimental investigation of the effect of fuselage-generated turbulence impinging on the leading edge of a swept wing on the boundary layer flow regime are presented. The possibilities of attenuating the turbulizing effect are studied. The criteria of laminar-turbulent transition are determined.     

Discontinuous solutions of an interacting three-dimensional hypersonic boundary layer

The paper is a mathematical study of the three-dimensional flow of viscous gas in a hypersonic boundary layer that develops along a flat wing whose leading edge has a step shape. The flow interacts with a flap on the wing set at a small angle. A linear solution to the problem is constructed under the assumption that the deflection angle of the flap is small and the difference between the length of the plates is of order unity. It is shown that an important part in the formation of the flow near and behind the flap may be played by the change in the pressure along the span of the wing due to the step shape of the leading edge. It is significant that although the pressure and displacement thickness are continuous functions of the transverse coordinate, the longitudinal and transverse components of the friction force have discontinuities.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 2, pp. 19–26, March–April, 1991.I thank V. V. Sychev and A. I. Ruban for suggesting the problem, for valuable advice, and assistance.     

Flow structure on a three-dimensional wing subjected to small amplitude perturbations

Small amplitude angular perturbations, of the order of one-half degree, can substantially modify the flow structure along a three-dimensional wing configuration, which is quantitatively characterized using a technique of high-image-density particle image velocimetry. Excitation at either the fundamental or the first subharmonic of the spanwise-averaged instability frequency of the separating shear layer from the stationary wing nearly eliminates the large-scale separation zone along the wing at high angle of attack. The physics of the flow is interpreted in terms of time-mean streamlines, vorticity and Reynolds stress, in conjunction with phase-averaged patterns of instantaneous vorticity. Distinctive vorticity patterns occur along the leading edge when the time-averaged separation zone is minimized.     

Stability of preseparation boundary layer on the leading edge of a thin airfoil

Numerous experiments on subsonic flow of gas past thin wing profiles (see the reviews [1, 2]) have shown that the flow near the leading edge of an airfoil is separationless only at angles of attack less than a certain critical value, which depends on the shape of the airfoil. If the angle of attack reaches the critical value, a closed region of recirculation flow of small extension is formed on the upper surface of the airfoil. Under ordinary flow conditions, the boundary layer on the leading edge of the airfoil remains laminar in the entire preseparation range of angles of attack. However, the appearance of the closed separation region is, as a rule, accompanied by transition from a laminar to a turbulent flow regime. Moreover, generation of turbulence is observed precisely in the flow separation region. The present paper is devoted to a study of the stability of the boundary layer on the leading edge of a thin airfoil in a flow of incompressible fluid. The case when the angle of attack of the airfoil relative to the oncoming flow differs little from the critical value is considered.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 6, pp. 55–63, November–December, 1982.     

Modeling the system of electrogasdynamic final-control elements

A system of electrogasdynamic final-control elements (plasma actuators) intended for increasing the stability of the boundary layer on a swept wing is considered. The actuators operate on the basis of dielectric barrier discharge. The physical model of the discharge in the air is formulated in the diffusion-drift approximation with account for three charged ionized-gas components, namely, electrons, positive nitrogen and oxygen ions, and negative oxygen ions. The boundary conditions on the dielectric surface are formulated with account for finite desorption and recombination rates of the charged particles. The numerical modeling for an actuator system of particular geometry shows a slight influence of the negative ions on the bulk force generated by each actuator. The main actuator parameters, such as the total longitudinal force and heat release, are shown to considerably depend on the dielectric permeabilities of insulation layers separating the external and internal electrodes. The expressions are derived that make it possible to estimate the gas velocity induced by the bulk force action of the dielectric barrier discharge on the gas flow.     

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 Heat Supply on the Stability of the Supersonic Swept Atiachment-Line Boundary Layer

The stability of a boundary layer with volume heat supply on the attachment line of a swept wing is investigated within the framework of the linear theory at supersonic inviscid-free-stream Mach numbers. The results of numerical calculations of the flow stability and neutral curves are presented for the flow on the leading edge of a swept wing with a swept angle χ=60° at various free-stream Mach numbers. The effect of volume heat supply on the characteristics of boundary layer stability on the attachment line is studied at a surface temperature equal to the temperature of the external inviscid flow. It is shown that in the case of a supersonic external inviscid flow volume heat supply may result in an increase in the critical Reynolds number and stabilization of disturbances corresponding to large wave numbers. For certain energy supply parameters the situation is reversed, the unstable disturbances corresponding to the main flow-instability zone are stabilized but another zone of flow-instability with small wave numbers and a significantly lower critical Reynolds number appears.     

Non-oscillating/Oscillating Shear Layer over a Large Deep Cavity at Low-Subsonic Speeds

The flow over a deep cavity at low subsonic velocity is considered in the present paper. The cavity length-to-depth aspect ratio is L/H = 0.2. Single hot-wire measurements characterized the incident turbulent boundary layer and show the influence of the cavity on the streamwise statistic components just downstream from the cavity. The streamwise mean and fluctuating velocity profiles are affected by the cavity. PIV measurements reveal the presence for ejection-like events responsible of local perturbations of the skewness and the flatness coefficients. Time-resolved PIV technic is also used to characterize phase properties of shear layer oscillating cycle. It is shown that for deep cavity with first Rossiter mode, only one vortical structure is formed at the cavity leading edge. Then, it grows while convecting downstream along the shear layer. A well-defined ejection process begins after the vortex impact near the cavity downstream corner. A cylinder device placed spanwisely near the cavity leading edge eliminates the resonance and highly modifies the behavior of the shear layer flow. In fact, the shear layer could be divided into upper and lower parts with different structure aspects.     

Flow along an attachment line controlled by a near-wall body force

The problems of the control of steady three-dimensional viscous incompressible flows modeling the gas flow in the vicinity of the attachment line on the leading edge of a swept wing are considered. It is assumed that the control is realized by means of the body force action approximating the action of a periodic sequence of plasma actuators mounted perpendicular to the leading edge. The corresponding boundary value problems for the system of Navier–Stokes equations are numerically solved using the Fourier method along the longitudinal coordinate and second-order difference approximation in the vertical coordinate. An important role played by the control-induced pressure gradient is shown.