Improving the Lift Properties of a Wing in Takeoff and Landing by Means of a Boundary Layer Control System Using Ejector-Type Actuators

A device making it possible to control of the flow past a wing at low flight speeds is proposed. The device comprises an ejector-type actuator that simultaneous sucks the boundary layer on the upper surface and blows a gas jet in the vicinity of the trailing edge. Based on the Reynolds-averaged Navier–Stokes equations, a mathematical model of an ejector-type actuator is developed and experimental studies are carried out.     

A numerical method for non-linear flow about a submerged hydrofoil

Summary A numerical method is presented for computing two-dimensional potential flow about a wing with a cusped trailing edge immersed beneath the free surface of a running stream of infinite depth. The full non-linear boundary conditions are retained at the free surface of the fluid, and the conditions on the hydrofoil are also stated exactly. The problem is solved numerically using integral-equation techniques combined with Newton's method. Surface profiles and the pressure distribution on the body are shown for different body geometries.     

Aerodynamic effects of flexibility in flapping wings

Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re ≈ 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small robotic insects and, to a limited extent, in understanding the aerodynamics of flapping insect wings.     

翼梢涡风洞研究中的摆动问题

采用激光粒子成像速度仪(PIV)对一矩形机翼(NACA0012)模型所产生的翼梢涡进行了风洞测试研究.测量位置为机翼近场尾迹,即x/c=3垂直于流动方向的截面,这里x为机翼后缘和测量截面之间的距离,c是机翼弦长.实验中基于弦长的雷诺数范围在3.4×104～26.6×104,通过分析所测得的涡量、切向速度和环量等,发现翼梢涡的摆动幅度与流过机翼上流体是否发生边界层分离有直接关系.     

叶栅内非定常不可压流动的数值模拟

基于SMAC(Simplified Marker and Cell)方法推导出直接求解二维非定常、不可压N-S方程的隐式数值方法.求解的基本方程是任意曲线坐标系中以逆变速度为变量的N-S方程和椭圆型的压力Poisson方程.采用该方法,对二维叶栅非定常分离流场进行了数值模拟,叶栅表面压力的计算结果与试验结果相比比较吻合,从而验证了这种方法的可靠性.同时对叶栅非定常流场的流场结构和流动机理做了初步的探讨.在均匀来流和定常边界条件下,叶栅内部流动表现出强烈的非定常性;在小冲角和高雷诺数时,叶栅尾部产生类似卡门涡街的周期性流动.     

Incompressible flow past oscillatory wings of low aspect ratio by the integral equations method

In the framework of the linearized perturbation theory, the pressure jump over the oscillating wing is the solution of a two‐dimensional integral equation. Performing an asymptotic expansion with respect to the aspect ratio and keeping the leading terms, we reduce the integral equation to a one‐dimensional one, obtaining a simplified method of solving the lifting surface integral equation for a class of thin wings of low aspect ratio with a straight trailing edge. The one‐dimensional integral equation is solved for the delta flat plate and the pressure coefficient field and the lift and moment coefficients are calculated. The range of validity of the new method is discussed in the final part of the paper. Copyright © 1999 John Wiley & Sons, Ltd.     

Singularities in BIEs for the Laplace equation; Joukowski trailing-edge conjecture revisited

The paper deals with trailing-edge issues connected with the analysis of three-dimensional incompressible quasi-potential flows (i.e. flows that are potential everywhere, except for a zero-thickness vortex layer, called the wake). Specifically, following the Joukowski conjecture of smooth flow at the trailing edge, all the trailing-edge conditions that are required to avoid singularities in the boundary integral representation for the velocity, in a quasi-potential incompressible flow around a wing, are identified. In particular, these include the Kondrat'ev and Oleinik singularity as well as the vortex-line and the edge-jet singularities of Epton. Also, following Mangler and Smith, the behavior of the wake geometry at the trailing edge is determined, using the Kutta condition of no pressure discontinuity at the trailing edge. Specific theoretical issues are addressed which include (1) the relationship between Joukowski conjecture and Kutta condition, and (2) identification of those trailing-edge conditions that are necessary to assure the uniqueness of the solution (as opposite to relationships that are automatically satisfied by the solution). Regarding the first issue, in the main body of the paper, the Joukowski conjecture and the Kutta condition are used as if they were independent assumptions; then, in Appendix A, it is shown that the Kutta condition need not be invoked as a separate assumption since it may be obtained as a consequence of the governing equations and of the Joukowski conjecture. In order to clarify the second issue, the theoretical analysis is coupled with a numerical one. In particular, the conditions necessary to insure uniqueness are inferred (not proven) through numerical experimentation: only the no-vortex-line condition appears to be necessary to insure uniqueness. This is accomplished by using a piecewise-cubic boundary-element method for quasi-potential flows that is an extension of a high-order formulation introduced by the authors and their collaborators (the order of the formulation is adequate to address all the theoretical trailing-edge conditions uncovered). The emphasis is on steady flows in simply connected regions; however, some issues related to unsteady flows in multiply connected regions are also examined. Finally, several open problems that require additional work are identified.     

Surface vorticity modeling of flow around airfoils

A surface vorticity method based on the new Kutta condition has been developed for solving the steady flow around an airfoil and unsteady separated flow past airfoil with a spoiler. In a discussion of the Kutta condition, it is argued that the appropriate Kutta condition is required to obtain a satisfactory solution. In this paper, two new methods of satisfying the Kutta condition incorporated in the surface vorticity method are described. The first method, which is based on results from flow visualization, is to introduce an additional control point at a short distance downstream of the trailing edge. In order to account for the fact that the velocity above the trailing edge is different from that below as in the real flow, the second method is to add a finite segment of vortex sheet downstream of the trailing edge. The present Kutta conditions are incorporated into the surface vorticity method and extend to solve unsteady flow around an airfoil with a spoiler. The computational results are in reasonable agreement with other computation as well as experiments.     

Uniqueness of the bounded flow solution in aerodynamics

We discuss uniqueness for steady incompressible inviscid flows past a body with a sharp trailing edge TE, with particular regard to multiconnected (toroidal) 3D wing configurations. Boundedness of the velocity field at TE is enforced by means of a singularity removal principal (Kutta condition). The resulting bounded flow solution is unique for 2D airfoils and 3D conventional wings. For toroidal bodies the flow depends on the available eigensolution which, however, has no direct influence on the lift. In this multiconnected case uniqueness of the bounded solution is shown to depend on the topology of the trailing edge.     

二维矩形柱体表面风压频域特性的雷诺数效应研究

在低紊流度的均匀流场中分别对截面宽厚比B/D(B为模型的顺风向宽度,D为迎风面厚度)为2、2.5、3、3.5和4的二维矩形柱体模型进行了测压试验,雷诺数(Re)的变化范围为1.1×10⁵~6.8×10⁵。分析了各模型表面的三个典型测点的脉动风压系数的功率谱特性,研究了各模型侧面的风压相关性。研究结果表明气流再附于模型侧面会影响模型侧面和背风面典型测点的风压系数功率谱,且B/D=4模型的三个典型测点的风压系数功率谱随雷诺数变化明显。B/D=2模型侧面的风压相关性较好,但B/D=2.5、3和3.5模型的侧面靠近前边缘的测点与气流再附区的测点的风压相关性较强,侧面中部区域的测点与侧面前后部区域的测点的风压相关性较差。B/D=4模型侧面的风压相关性受雷诺数影响明显,在雷诺数Re=6.8×10⁵时出现了明显的突变。     

排式双翼布局低雷诺数气动特性计算研究

作为一种新型的气动布局形式,排式布局对低雷诺数流动具有较高的气动效率,适用于柔性可充气飞行器,比如充气式飞机或是高空飞艇。但是,由于前、后翼之间强烈的气动干扰现象,目前对此类布局的气动特性认识还十分有限。为了充分理解这种布局的气动特点,在前期风洞试验的基础上,开展了数值模拟工作,详细地研究了低雷诺数情况下翼型厚度、表面波纹状外形及后翼偏转角度等几何因素对此类飞行器气动特性的影响规律。计算结果表明,在计算的迎角范围内,排式布局能通过前后翼之间的气动干扰延缓或抑制机翼后缘处的流动分离,从而提高整体气动效率,因此排式布局在未来很适合应用于小型无人机或是飞艇等可充气式飞行器构型上。     

某跨音速高压涡轮非定常流动数值模拟研究

通过求解三维非定常雷诺平均N-S方程模拟某跨音速高压涡轮非定常流场,研究涡轮内非定常流动特征。通过对静子尾迹及静子尾缘激波和转子叶排之间的相互干涉过程进行详细分析,发现定常/非定常模拟方法获得的涡轮总体性能参数基本一致但流场存在较大差异。静子尾迹是导致涡轮流场非定常性的重要因素之一：在转子叶栅通道中部和下部,静子尾迹和转子叶片附面层及下通道涡发生明显干涉,并导致通道中下部损失周期性波动幅度较大,此外尾迹和下通道涡间的干涉作用在转子尾缘处诱导出高频脱落涡。静子尾缘激波也是导致涡轮流场非定常性的原因之一,激波和转子叶片作用形成复杂的波系结构,对涡轮流场影响显著：一方面激波/附面层干涉导致转子和静子的吸力面产生周期性变化的高温区域;另一方面激波撞击叶片导致叶片表面的气流在激波后出现分离,对转子静压分布产生影响,使得转子叶片表面载荷出现明显的非定常性,进而导致涡轮输出功的周期性波动十分剧烈。     

Large‐eddy simulation of separated leading‐edge flow in general co‐ordinates

An incompressible separated transitional boundary‐layer flow on a flat plate with a semi‐circular leading edge has been simulated and a very good agreement with the experimental data has been obtained, demonstrating how this technique may be applied even when finite difference formulae are used in the periodic direction. The entire transition process has been elucidated and vortical structures have been identified at different stages during the transition process. Efficient numerical methods for the large‐eddy simulation (LES) of turbulent flows in complex geometry are developed. The methods used are described in detail: body‐fitted co‐ordinates with the contravariant velocity components of the general Navier–Stokes equations discretized on a staggered mesh with a dynamic subgrid‐scale model in general co‐ordinates. The main source of computational expense in simulations for incompressible flows is due to the solution of a Poisson equation for pressure. This is especially true for flows in complex geometry. Fourier techniques can be employed to speed up the pressure solution significantly for a flow which is periodic in one dimension. With simple conditions fulfilled, it is possible to Fourier transform a discrete elliptic equation such as the Poisson equation for the pressure field, decomposing the problem into a set of two‐dimensional problems of similar type (Poisson‐like). Even when a complex geometry and body‐fitted curvilinear co‐ordinates are used in the other two dimensions, as in the present case, the resulting Fourier‐transformed 2D problems are much more efficiently solved than the 3D problem by iterative means. Copyright © 2000 John Wiley & Sons, Ltd.     

A vortex method for separated flow around an airfoil with a detached spoiler

A discrete vortex method based on no-slip condition is developed for simulating unsteady separated flows around an airfoil with a detached spoiler. For flow separated at each sharp edge, such as the spoiler tips and the trailing edge of the airfoil, a vortex sheet is used to feed discrete vortices at each time step. The length, inclination and strength of each sheet is determined by the continuity equation, the momentum principle and a Kutta pressure condition such that the flow, net force and pressure difference across the vortex sheet are all zero. The separation on the airfoil upper surface is simulated by discrete vortices shed from a fixed separation point. The flow patterns behind a detached spoiler at different time steps are obtained and compared with those of the conventional spoiler. Reasonable agreements are found between the predicted pressure distributions and experimental measurements. The computed results show that base-venting changes the flow field around the spoiler and reduces the adverse effect in lift experienced by the airfoil when the spoiler undergoes a rapid deployment.     

An inverse method for transonic wing design

A fast, efficient and reliable code for wing design has been developed at INTA coupling a residual‐correction method by Bauer, McFadden and Garabedian to an inviscid solver for wing analysis in transonic flow. Smoothing procedures, including Bèzier cubic splines, are used to avoid irregularities of the wing surface, as well as the twist distribution. A modified version of FLO22 code is used as the flow solver. The original code has been adapted to improve its accuracy. Some results are presented, showing the reliability of the code. The redesign of a wing in transonic flow—which was used as test case in LARA project of BRITE‐EURAM II during 1993–94—is presented with promising results. Copyright © 1999 John Wiley & Sons, Ltd.     

后缘小翼智能旋翼减振效果影响因素分析

建立了带后缘小翼智能旋翼气动弹性载荷计算模型及减振优化分析方法。模型考虑刚体后缘小翼的气动力与惯性力对弹性桨叶系统的影响,使用粘性涡粒子法结合翼型查表法计算旋翼气动载荷,采用力积分法计算桨叶与桨毂载荷,构造了包含桨叶根部扭转及桨毂振动载荷为目标函数的优化问题,基于最速下降-黄金分割组合优化算法寻找最佳小翼偏转规律。研究发现,建立的后缘小翼载荷控制方法有效,可降低振动目标函数70%。桨叶的弹性扭转使后缘小翼能有效实施减振,但弹性扭转对小翼气动力矩的放大作用使减振时通常伴随着桨叶扭转载荷增大的现象。     

机翼屏蔽对航空发动机噪声预测的影响分析

随着对飞机适航噪声的要求更加严格,利用机翼屏蔽效应对航空发动机进行噪声控制已成为一项有效策略与研究方向.基于惠更斯-菲涅尔原理,利用飞机噪声性能(Aircraft Noise and Performance,ANP)数据库计算飞机起飞航迹并在其上建立噪声源,声源位置作为噪声屏蔽计算坐标输入屏蔽效应算法,应用Matlab...     

An adaptive fixed mesh method for modelling and simulating of unsteady two‐phase flows with sharp physical interfaces

We present in this paper a new computational method for simulation of two‐phase flow problems with moving boundaries and sharp physical interfaces. An adaptive interface‐capturing technique (ICT) of the Eulerian type is developed for capturing the motion of the interfaces (free surfaces) in an unsteady flow state. The adaptive method is mainly based on the relative boundary conditions of the zero pressure head, at which the interface is corresponding to a free surface boundary. The definition of the free surface boundary condition is used as a marker for identifying the position of the interface (free surface) in the two‐phase flow problems. An initial‐value‐problem (IVP) partial differential equation (PDE) is derived from the dynamic conditions of the interface, and it is designed to govern the motion of the interface in time. In this adaptive technique, the Navier–Stokes equations written for two incompressible fluids together with the IVP are solved numerically over the flow domain. An adaptive mass conservation algorithm is constructed to govern the continuum of the fluid. The finite element method (FEM) is used for the spatial discretization and a fully coupled implicit time integration method is applied for the advancement in time. FE‐stabilization techniques are added to the standard formulation of the discretization, which possess good stability and accuracy properties for the numerical solution. The adaptive technique is tested in simulation of some numerical examples. With the test problems presented here, we demonstrated that the adaptive technique is a simple tool for modelling and computation of complex motion of sharp physical interfaces in convection–advection‐dominated flow problems. We also demonstrated that the IVP and the evolution of the interface function are coupled explicitly and implicitly to the system of the computed unknowns in the flow domain. Copyright © 2005 John Wiley & Sons, Ltd.     

Computational investigation of cicada aerodynamics in forward flight

Free forward flight of cicadas is investigated through high-speed photogrammetry, three-dimensional surface reconstruction and computational fluid dynamics simulations. We report two new vortices generated by the cicada''s wide body. One is the thorax-generated vortex, which helps the downwash flow, indicating a new phenomenon of lift enhancement. Another is the cicada posterior body vortex, which entangles with the vortex ring composed of wing tip, trailing edge and wing root vortices. Some other vortex features include: independently developed left- and right-hand side leading edge vortex (LEV), dual-core LEV structure at the mid-wing region and near-wake two-vortex-ring structure. In the cicada forward flight, approximately 79% of the total lift is generated during the downstroke. Cicada wings experience drag in the downstroke, and generate thrust during the upstroke. Energetics study shows that the cicada in free forward flight consumes much more power in the downstroke than in the upstroke, to provide enough lift to support the weight and to overcome drag to move forward.