Generation of Three-Dimensional Disturbances in a Boundary Layer on Strong Interaction with a Hypersonic Flow

Laminar boundary layer flow over an infinite-span, finite-length flat plate is investigated in the regime of strong interaction with a hypersonic gas flow. Under the assumption that an additional condition dependent on the transverse coordinate can be imposed on the trailing edge of the plate the flow functions are expanded in power series in the vicinity of the leading edge. It is shown that these expansions include an indefinite function dependent on the transverse coordinate. The corresponding boundary value problems are formulated and solved and the eigenvalues are determined. It is established that in this case the two-dimensional boundary layer can rearrange itself into a three-dimensional boundary layer.     

Controlling the disturbances in a hypersonic flat-plate shock layer by unsteady action from the surface

The results of a wind-tunnel experiment on the joint action of periodic acoustic fast-mode disturbances of the outer flow and disturbances generated at the leading edge of a plate on the hypersonic (M_∞ = 21) viscous shock layer on the plate are presented. The possibility of positively controlling the intensity of density fluctuations in the plate shock layer by means of disturbances introduced from the leading edge is shown. Direct numerical simulation of the suppression (enhancement) of disturbances under the simultaneous action on the shock layer of the two-dimensional fast-mode acoustic waves in the outer flow and the source of two-dimensional suction/injection disturbances near the leading edge of the plate is performed under the experimental conditions. The experimental and calculated results are shown to be in good agreement.     

Receptivity of a flat plate boundary layer to a free stream axial vortex

A single free stream axial vortex of controlled strength and position was used to investigate a vortical receptivity of Blasius boundary layer. Excited boundary-layer disturbances were dominated by streamwise velocity perturbations, that grew downstream essentially linearly with the streamwise coordinate. It was shown that the disturbance characteristics are in agreement with data of previous experiments performed under natural and control conditions concerning the ‘by-pass’ transition initiated at high free stream disturbance levels. It was proved that the role of the leading edge in the receptivity process and disturbance growth under consideration is not dominant.     

Stabilization of flow over an attachment line using a near-wall body force

The problem of the control of steady flow in the vicinity of the attachment line on the leading edge of a swept wing is considered. It is assumed that the control is realized by means of near-wall body force action for the purpose of increasing the flow stability. The stability problem is solved under the assumption of the streamwise homogeneity of the undisturbed three-dimensional flow. The dependence of the stabilizing effect on the action amplitude and the width of the region, where the near-wall force is applied, is studied.     

On the Optimal Shapes of Waveriders Constructed on Plane Shocks

Formulations of variational problems on maximum lift-drag ratio lifting shapes are considered for different sets of isoperimetric conditions. The problem with a differential constraint setting a lower limit on the local slope of the leading edge of the waverider and simulating either the maximum heat flux to the leading edge or the contribution of the force acting on the leading edge to a particular component of the aerodynamic force is considered. Solutions of the problem of the optimal shape of a waverider constructed on plane shocks are derived for given lift coefficient and specific volume, both with and without constraints on the waverider dimensions.     

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.      

Singular perturbations in free convection

A second order solution to the problem of free convection from a vertical surface has been obtained using the method of perturbations. It has been shown that the perturbation solution gives divergent solutions at the leading edge, but is otherwise satisfactory. Further, using the first order outer solution, a coordinate system, optimal in the sense ofKaplun, has been obtained. It is shown that this coordinate system presents a free convection flow field which is more satisfactory than that given by the classical Pohlhausen solution.     

Investigation of flows past a plate in the presence of exoergic processes driven by the interaction between a nonequilibrium electric discharge and a propane-air mixture

Supersonic flow (M = 2) past a plate along which propane is injected is investigated within the framework of the solution of problems of combustion initiation and stabilization at low static temperatures and pressures in the presence of a nonequilibrium discharge and metal and dielectric interceptors mounted on the plate surface. The experiments show that two zones with exoergic reactions develop when a metal interceptor is mounted on the plate. One zone is located ahead of the leading separation zone and the other above and behind the interceptor edge, its boundary partially penetrating into the supersonic flow region. Using modern spectroscopic methods, the radiation intensity distributions of a series of plasmochemical reaction products are obtained in the neighborhood of the plate ahead of the interceptor, behind it, and above its edge. It is found that the fuel is intensively converted under the action of the discharge with the occurrence of a series of free radicals, atomic hydrogen and oxygen which are themselves chemically active.     

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.     

Behavior of a Semi-Infinite Ice Cover Under a Uniformly Moving Load

This paper consideres the behavior of a semi-infinite ice cover on the surface of an ideal incompressible fluid of finite depth under the action of a load moving with constant velocity along the edge of the cover at some distance from it. The ice cover is modeled by a thin elastic plate of constant thickness. In a moving coordinate system, the deflection of the plate is assumed to be steady. An analytic solution of the problem is obtained using the Wiener–Hopf technique. The wave forces, the deflection of the plate, and the elevation of the free surface of the fluid at different velocities of the load are investigated.     

Flapping dynamics of a flexible propulsor near ground

The flapping motion of a flexible propulsor near the ground was simulated using the immersed boundary method. The hydrodynamic benefits of the propulsor near the ground were explored by varying the heaving frequency (St) of the leading edge of the flexible propulsor. Propul-sion near the ground had some advantages in generating thrust and propelling faster than propulsion away from the ground. The mode analysis and flapping amplitude along the Lagrangian coordinate were examined to analyze the kine-matics as a function of the ground proximity (d)and St. The trailing edge amplitude (atail)and the net thrust (Fx)were influenced by St of the flexible propulsor. The vortical structures in the wake were analyzed for different flapping conditions.     

Adaptive Feedback Linearization for the Control of a Typical Wing Section with Structural Nonlinearity

Earlier results by the authors showed constructions of Lie algebraic, partial feedback linearizing control methods for pitch and plunge primary control utilizing a single trailing edge actuator. In addition, a globally stable nonlinear adaptive control method was derived for a structurally nonlinear wing section with both a leading and trailing edge actuator. However, the global stability result described in a previous paper by the authors, while highly desirable, relied on the fact that the leading and trailing edge actuators rendered the system exactly feedback linearizable via Lie algebraic methods. In this paper, the authors derive an adaptive, nonlinear feedback control methodology for a structurally nonlinear typical wing section. The technique is advantageous in that the adaptive control is derived utilizing an explicit parameterization of the structural nonlinearity and a partial feedback linearizing control that is parametrically dependent is defined via Lie algebraic methods. The closed loop stability of the system is guaranteed to be stable via application of La Salle's invariance principle.     

Two contact problems for a wedge with a symmetric cut on the edge

In this paper, the solutions of two contact problems for a wedge with a symmetric cut on the edge are presented. First, special approximation methods and orthogonal polynomials are used to solve the auxiliary problem on the action of a lumped force on the cut edge. The obtained solutions are compared with known solutions in some special cases.     

高超声速飞行器气动防热新概念研究

传统乘波构型的高超声速飞行器尖锐的前缘存在严重的气动加热问题,而简单的前缘钝化气动防热方法由于造成很大的升阻比损失,难以发挥实质性作用. 引入``人工钝前缘(ABLE)'概念,拟以一种新的思路解决这一矛盾. 通过定义ABLE构型的外形参数,并采用CFD数值计算方法研究了各参数对气动力和气动热特性的影响规律,在流场分析的基础上进行了外形优化,最终得到令人满意的新型高超声速飞行器头部外形,总结了运用ABLE概念进行气动防热的相关设计原则和规律.      

Computation of two‐dimensional flows past ram‐air parachutes

Computational results for flow past a two‐dimensional model of a ram‐air parachute with leading edge cut are presented. Both laminar (Re=10⁴) and turbulent (Re=10⁶) flows are computed. A well‐proven stabilized finite element method (FEM), which has been applied to various flow problems earlier, is utilized to solve the incompressible Navier–Stokes equations in the primitive variables formulation. The Baldwin–Lomax model is employed for turbulence closure. Turbulent flow computations past a Clarck‐Y airfoil without a leading edge cut, for α=7.5°, result in an attached flow. The leading edge cut causes the flow to become unsteady and leads to a significant loss in lift and an increase in drag. The flow inside the parafoil cell remains almost stagnant, resulting in a high value of pressure, which is responsible for giving the parafoil its shape. The value of the lift‐to‐drag ratio obtained with the present computations is in good agreement with those reported in the literature. The effect of the size and location of the leading edge cut is studied. It is found that the flow on the upper surface of the parafoil is fairly insensitive to the configuration of the cut. However, the flow quality on the lower surface improves as the leading edge cut becomes smaller. The lift‐to‐drag ratio for various configurations of the leading edge cut varies between 3.4 and 5.8. It is observed that even though the time histories of the aerodynamic coefficients from the laminar and turbulent flow computations are quite different, their time‐averaged values are quite similar. Copyright © 2001 John Wiley & Sons, Ltd.     

Biquaternion solution of the kinematic control problem for the motion of a rigid body and its application to the solution of inverse problems of robot-manipulator kinematics

The problem of reducing the body-attached coordinate system to the reference (programmed) coordinate system moving relative to the fixed coordinate system with a given instantaneous velocity screw along a given trajectory is considered in the kinematic statement. The biquaternion kinematic equations of motion of a rigid body in normalized and unnormalized finite displacement biquaternions are used as the mathematical model of motion, and the dual orthogonal projections of the instantaneous velocity screw of the body motion onto the body coordinate axes are used as the control. Various types of correction (stabilization), which are biquaternion analogs of position and integral corrections, are proposed. It is shown that the linear (obtained without linearization) and stationary biquaternion error equations that are invariant under any chosen programmed motion of the reference coordinate system can be obtained for the proposed types of correction and the use of unnormalized finite displacement biquaternions and four-dimensional dual controls allows one to construct globally regular control laws. The general solution of the error equation is constructed, and conditions for asymptotic stability of the programmed motion are obtained. The constructed theory of kinematic control of motion is used to solve inverse problems of robot-manipulator kinematics. The control problem under study is a generalization of the kinematic problem [1, 2] of reducing the body-attached coordinate system to the reference coordinate system rotating at a given (programmed) absolute angular velocity, and the presentedmethod for solving inverse problems of robotmanipulator kinematics is a development of the method proposed in [3–5].     

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。     

Experimental study of vortex–structure interaction noise radiated from rod–airfoil configurations

Vortex–structure interaction noise radiated from an airfoil embedded in the wake of a rod is investigated experimentally in an anechoic wind tunnel by means of a phased microphone array for acoustic tests and particle image velocimetry (PIV) for the flow field measurements. The rod–airfoil configuration is varied by changing the rod diameter (D), adjusting the cross-stream position (Y) of the rod and the streamwise gap (L) between the rod and the airfoil leading edge. Two noise control concepts, including “air blowing” on the upstream rod and a soft-vane leading edge on the airfoil, are applied to control the vortex–structure interaction noise. The motivation behind this study is to investigate the effects of the three parameters on the characteristics of the radiated noise and then explore the influences of the noise control concepts. Both the vortex–structure interaction noise and the rod vortex shedding tonal noise are analysed. The acoustic test results show that both the position and magnitude of the dominant noise source of the rod–airfoil model are highly dependent on the parameters considered. In the case where the vortex–structure interaction noise is dominant, the application of the air blowing and the soft vane can effectively attenuate the interaction noise. Flow field measurements suggest that the intensity of the vortex–structure interaction and the flow impingement on the airfoil leading edge are suppressed by the control methods, giving a reduction in noise.     

Leading edge slot blowing on an open cavity in supersonic flow

Supersonic flow over an open cavity can create intense pressure loads on the surfaces within the cavity. In order to combat these loads, the development of a control scheme to reduce them is becoming increasingly important for many engineering applications. The present study implements steady leading edge blowing through various configurations of spanwise-aligned rectangular leading edge slots. The effects of this control on the flow field were examined to determine the suppression mechanisms exploited by the leading edge blowing. The cavity studied here had a length-to-depth ratio of 6 and was placed in a freestream flow with a Mach number of 1.4. Actuators with one continuous slot and three and five segmented slots spanning the width of the cavity were installed at the leading edge. Surface pressure reductions of nearly 45% were achieved on the aft wall of the cavity using the 5-slot configuration. Velocity field measurements acquired through 2-component (streamwise-aligned measurement plane) and 3-component stereoscopic (cross-stream-aligned measurement plane) particle image velocimetry revealed the presence of streamwise-aligned vortices created by the segmented slots. These act to significantly alter the shear layer formed at the mouth of the cavity creating highly three-dimensional flow field features.     

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.