A modified discrete-vortex method algorithm with shedding criterion for aerodynamic coefficients prediction at high angle of attack

Low-order methods require less computing power than classical computational fluid dynamics and can be implemented on a laptop computer, which is needed for engineering tasks. Discrete vortex methods are such low order methods that can describe the unsteady separated flow around an airfoil. After a presentation of the leading edge suction parameter discrete vortex method, a modified algorithm is proposed, in order to reduce the computing cost, and compared with the previous one. Several reference unsteady airfoil motions are discussed in terms of gain in the computation time with comparisons between the previous scheme and the present one. The accuracy of the new method is demonstrated through aerodynamic coefficients. The application of the present discrete vortex method to a transient pitching motion of an airfoil is also presented, in order to understand the leading edge vortex formation, and its implication in terms of lift and drag coefficients. The method is not limited to unsteady or transient motions but can also simulate the flow around a constant angle of attack airfoil. In that case, an original method of fast summation of the vortices located far away from the airfoil, allows a linear dependence of the computation time versus the number of vortices shed, which is a great improvement over the quadratic dependence observed in the classical discrete vortex methods. The development of the aerodynamic coefficients with angle of attack, from values ranging between −10° and 90°, is obtained for a purely two-dimensional flow. In particular, the shape of the lift coefficient of the airfoil in the fully detached flow region is established. Comparisons with relevant experimental or computational fluid dynamics data are discussed in order to grasp the influence of upstream turbulence level and three-dimensional effects in the measured data in the fully detached flow region.     

风力机叶型增升及叶尖流动控制研究

研究了两种改善风力机叶型气动性能的流动控制技术,分别对风力机专用S809翼型和较大升阻比的FX 60-100翼型进行应用研究．首先,通过在叶型前缘加装流动偏转器,研究流动偏转器对叶型流动分离的控制效果．并采用多岛基因算法,对流动偏转器进行多参数优化．结果表明:流动偏转器可以有效控制叶型的失速特性,推迟失速攻角和增加升力;基因优化算法能更大地提升流动偏转器的控制效果．其次,基于对风力机叶尖旋涡和尾涡特征以及叶片表面压力分布的分析,在叶片尖部加装不同倾斜角的旋涡扩散器控制叶尖涡．结果表明:涡扩散器能够提高叶尖涡涡核的总压,削弱其旋涡强度,使风力机尾流旋涡耗散更快,从而可以减小噪声,提高叶片效率．     

翼型绕流干涉噪声的实验与数值研究

来流湍流干扰噪声在风力机叶片气动总噪声级中占有重要地位．选取圆柱/翼型干涉模型从实验和数值两方面研究此类干涉发声现象．实验中通过对翼型表面非定常载荷的测量,重点研究了圆柱位置和翼型攻角的影响,选取的翼型包括两个NACA系列翼型（NACA0012和NACA0018）和两个风力机翼型（s809和s825）,同时利用PIV（particle image velocimetry）技术对低攻角状态下翼型的前缘流场进行了研究．实验结果表明翼型表面非定常压力与圆柱涡脱落存在一定相关性．与此同时采用非定常Reynolds平均(URANS)方法对圆柱/NACA0012翼型的干涉流场进行了非定常数值模拟,并将得到的翼型表面压力频谱与实验结果进行了对比．     

An insight into the rotational stall delay

Blade element momentum (BEM) theory which is based on the two-dimensional (2D) aerodynamic properties of airfoil blade element is the most common computational engineering method for the prediction of loads and power curves of wind turbines. Although most BEM models yield acceptable results for high tip-speed ratios where the local angles of attack are small, no generally accepted model exists up to date that consistently predicts the loads and power in stall regime for stall-controlled turbines. Understanding of the stall delay phenomenon on wind turbines remains, to this day, incomplete. The lack of a conceptual model for the complex three-dimensional (3D) flow field on the rotor blade, where stall is begins, how it progresses and where stall is practically terminated, has hindered the finding of a unanimously accepted solution. The paper aims at giving a better understanding of the delayed stall events and a reasonably simple correction model that complements the 2D airfoil characteristics used to a BEM method. (© 2012 Wiley-VCH Verlag GmbH & Co. KGaA, Weinheim)     

Numerical bifurcation analysis of static stall of airfoil and dynamic stall under unsteady perturbation

By the finite element method combined with Arbitrary-Lagrangian-Eulerian (ALE) frame and explicit Characteristic Based Split Scheme (CBS), the complex flows around stationary and sinusoidal pitching airfoil are studied numerically. In particular, the static and dynamic stalls are analyzed in detail, and the natures of the static stall of NACA0012 airfoil are given from viewpoint of bifurcations. Following the bifurcation in Map, the static stall is proved to be the result from saddle-node bifurcation which involves both the hysteresis and jumping phenomena, by introducing a Map and its Floquet multiplier, which is constructed in the numerical simulation of flow field and related to the lift of the airfoil. Further, because the saddle-node bifurcation is sensitive to imperfection or perturbation, the airfoil is then subjected to a perturbation which is a kind of sinusoidal pitching oscillation, and the flow structure and aerodynamic performance are studied numerically. The results show that the large-scale flow separation at the static stall on the airfoil surface can be removed or delayed feasibly, and the ensuing lift could be enhanced significantly and also the stalling incidence could be delayed effectively. As a conclusion, it can be drawn that the proper external excitation can be considered as a powerful control strategy for the stall. As an unsteady aerodynamic behavior of high angle of attack, the dynamic stall can be investigated from viewpoint of nonlinear dynamics, and there exists a rich variety of nonlinear phenomena, which are related to the lift enhancement and drag reduction.     

Simulation of Aerodynamic Performances of Flexible Flapping Wing Airfoils北大核心CSCD

与固定翼相比,在低速、小Reynolds数条件下,扑翼飞行具有显著的气动性能优势,受到越来越多的重视。然而,目前对扑翼翼型的研究以刚性翼型为主,对柔性翼型气动性能认识还不清楚。该文建立了柔性椭圆翼型的流固耦合仿真模型,分析了不同风速、迎角下柔性椭圆翼型的周围流场、变形以及气动性能。仿真结果表明,较刚性翼型,柔性翼型延缓了尾涡脱落时间,有效降低升力扰动振荡频率;柔性翼型显著抑制了尾流流场的扰动,降低升力扰动振荡幅值,合适的弹性模量翼型使得扰动振荡完全消除。研究结果可为软飞行器气动设计提供参考。     

Leading edge separation: a comparison between interaction theory and Navier–Stokes computations

We consider the two‐dimensional high/moderate Reynolds‐number flow around a slender airfoil at small angles of attack. Special emphasis is placed on leading edge separation and the associated breakdown of the steady solution of the governing equations. A comparison between the predictions of the viscous/inviscid interaction theory of marginal separation, see e.g. [3], and the results of Navier–Stokes computations with DLR's ‘TAU’–code [2] shows excellent quantitative agreement concerning the location, size of the separation bubble and even the limiting maximum value of the angle of attack, up to which solutions by means of the particular approaches can be found.     

Quasisolution of an inverse boundary-value problem in aerohydrodynamics with a limited maximum velocity on a contour in a range of angles of attack

In this paper we develop a method for constructing a quasisolution to the inverse boundary value problem in aerohydrodynamics with a limited maximum velocity on an airfoil surface in a range of angles of attack. We reduce this problem to the minimization of a quadratic functional subject to constraints in the form of equalities and inequalities.     

A stall-delay model for rotating blades

The paper is aimed at describing a fundamental phenomenon: the effect of rotation on the inboard blade boundary layer on a wind turbine. The three-dimensional incompressible steady momentum integral boundary-layer equations are used to predict the boundary-layer growth and limiting streamline angles on the blade surface for both attached and separating flow. The chordwise skin friction coefficient is used to identify boundary layer separation and shear layer reattachment locations. The nature of flow near the axis of rotation is discussed and the physical mechanism associated with 3-D and rotational effects is identified. A semi-empirical correction law for the lift coefficient based on 2-D airfoil data is established. (© 2008 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim)     

On Transonic Marginal Separation

Laminar boundary‐layer separation near the leading edge of a thin airfoil is one of the principal factors that limits the lift force acting on the airfoil. Marginal separation, in particular, denotes the onset of separation which is accompanied by the formation of a short separation bubble. Using asymptotic analysis this effect is studied in the limit of high Reynolds number and for transonic external flow conditions. It is assumed that the fluid under consideration is a perfect gas and the airfoil surface is taken to be thermally insulated. Results to be presented include the analytical investigation of the emerging three layer structure, the associated transonic far field and the calculation of representative wall shear stress distributions. (© 2004 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)     

基于线性涡模型的风洞孔壁干扰特性分析

发展了一种适用于二元翼型试验洞壁干扰特性的评估和修正方法.基于Prandtl-Glauert速度势方程和布置在模型及洞壁表面的线性涡,采用迭代方法计算了风洞孔壁对翼型表面压力分布特性的影响,分析了不同孔壁透气特性参数的影响规律和量值,利用与国外参考结果及风洞试验结果的对比确定了该方法的准确性.结果表明,孔壁对翼型绕流的影响主要反映在上翼面吸力峰和最大厚度位置之间,使压力系数减小,积分后的升力系数降低,且随着孔壁透气特性参数的增大,洞壁干扰由实壁特性向开口特性发展,洞壁干扰、影响量急剧增大.与传统方法相比,该方法计算快速,结果可靠,同时具备试验前评估的能力,可用于亚临界范围内翼型表面压力的快速估算,以及翼型试验的洞壁干扰修正.     

Pressure and flow in two dimensional numerical bifurcation models

The finite element method has been used to solve the Navier-Strokes equations for steady flow conditions in bifurcations. The results are presented as pressure, velocity and streamline plots at different Reynolds number. The three bifurcations considered have rigid walls and bifurcation angles of 0°, 20° and 180°. For the bifurcation with branch angles 0° and 20° there is flow separation along the inner wall of the outlet branches and large spatial pressure variations, these phenomena being more pronounced at the higher Reynolds numbers. For the bifurcation with a branch angle of 180° the high pressure gradients occured at the outer corner and for the high Reynolds number a vortex formation developed downstream of this corner.     

On higher order effects in marginally separated flows

If the angle of attack α of a slender airfoil reaches a critical value α_s flow separation is known to occur at the upper s surface. Further increase of α initially leads to the formation of a short laminar separation bubble which has an extremely weak influence on the external flow field – a phenomenon known as marginal separation – but then rather rapidly causes a severe change of the flow behaviour, leading to leading edge stall. According to the asymptotic theory of marginal separation holding in the limit of large Reynolds numbers Re, the flow in the neighbourhood of the separation bubble is governed by an integro-differential equation. This so-called interaction equation contains a single controlling parameter which relates the angle of attack to the Reynolds number, with a value Γ_s corresponding to α_s. Some recent results concerning higher order s s corrections to this theory and their effect on the stability of steady solutions will be presented. (© 2011 Wiley-VCH Verlag GmbH & Co. KGaA, Weinheim)     

Numerical study of helicopter blade–vortex mechanism of interaction using the potential flow theory

The blade–vortex interaction (BVI) phenomenon plays a key role in the rotorcraft aerodynamics. Numerical investigations of BVI using classical CFD approaches are computationally expensive. In the present research we propose a numerical approach, based on the potential flow theory, for the numerical investigation of helicopter blade–vortex mechanism of interaction. This approach overcomes the computational expenses posed by the CFD techniques. The influence of vertical miss distance, angle of attack, airfoil camber, and vortex strength on the helicopter blade–vortex mechanism of interaction is subject of investigation. The study reveals that the magnitude of the aerodynamic coefficients decreases with the increase of vertical miss distance and angle of attack, and the decrease of vortex strength and core size.     

Near-wall gas-droplet flows with liquid-film formation

A self-consistent asymptotic theory of a near-wall aerodisperse flow with the formation of a liquid film due to inertial deposition of liquid droplets on the body surface is constructed. With the reference to the examples of low-speed mist flows near plane and curvilinear solid surfaces, asymptotically different situations are analyzed. The two-fluid equations of the mist flow are solved simultaneously with the equations of the liquid film, on the outer edge of which the distributed mass, momentum, and energy fluxes (found from the solution of the outer problem) are specified. It is shown that depending on the values of governing parameters the flow in the film should be described by either boundary-layer or creeping-flow equations in the layer with a priori unknown thickness. For the film on the frontal surface of a blunt body, the film thickness and the friction and heat transfer coefficients are found numerically. The conditions in which the presence of the film significantly reduces the heat fluxes are determined and the possibility of the realization of steady-state flow regimes in the film is demonstrated. (© 2005 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim)     

Use of active control for elimination of flow separation with estimation of energy costs

Problems were posed and solved concerning the aerodynamic computation of the flow past an airfoil with an active boundary layer control device used to prevent flow separation. A moving wall, suction, or tangential blowing in the boundary layer was used as a flow control device. The turbulent boundary layer was computed by directly solving the boundary layer equations using an implicit difference scheme with adaptive grid generation and the determination of the computational domain size. A software code was developed, and numerical simulations were performed taking into account the energy costs related to the flow control device. The numerical results showed that the active flow control devices can be used to prevent flow separation.     

Frequency analysis of carbon and silicon nanosheet with surface effects

Silicon-based microelectromechanical system (MEMS) and nanoelectromechanical systems (NEMS) have been used to design and fabricate sensitive sensors and actuators. Recent research trends show that graphene and carbon nanotubes (CNTs) have been used to change the surface properties of silicon-based MEMS and NEMS to improve different mechanical, optical and electrical properties of silicon-based composites. In this paper, we focus on analyzing the vibrational characteristics of silicon-based devices when the surface of silicon is coated with single-layer graphene and horizontally aligned carbon nanotubes (HACNTs). To perform the analysis, we use multi-scale finite element approach for developing graphene–silicon nanocomposites (GSNCs) and carbon nanotube-silicon nanocomposites (CSNC) composites in which interface layer of silicon with graphene or CNT is modeled using bonded contact element. Subsequently, we performed modal analysis to find the first transverse mode frequency of GSNC and CSNC composites for beam with smaller as well as longer lengths. The numerical model is compared with classical beam theory with and without surface effect. For GSNCs composites, we take a fixed-free case with lengths in the range of (20 Å–120 Å) and (400 Å–2000 Å), respectively. For CSNC composites, CNT diameter is varied from (5 Å–30 Å) for single walled nanotube. Subsequently, we analyze the influence of HACNTs-on-silicon on its vibrational characteristics. The analysis presented in the paper demonstrate that GSNCs offer a higher bending stiffness compared to single layer graphene (SLGs) and isolated silicon nanosheet which lead to higher natural frequency. A similar trend is found in the case of HACNTs on silicon NS when the number of tubes increases.