Calculation of nonlinear aerodynamic characteristics of a wing using a 3‐D panel method

The nonlinear aerodynamic characteristic of a wing is investigated using the frequency‐domain panel method. To calculate the nonlinear aerodynamic characteristics of a three‐dimensional wing, the iterative decambering approach is introduced into the frequency‐domain panel method. The decambering approach uses the known nonlinear aerodynamic characteristic of airfoil and calculates two‐variable decambering function to take into consideration the boundary‐layer separation effects for the each section of the wing. The multidimensional Newton iteration is used to account for the coupling between the different sections of wing. The nonlinear aerodynamic analyses for a rectangular wing, a tapered wing, and a wing with the control surface are performed. Present results are given with experiments and other numerical results. Computed results are in good agreement with other data. This method can be used for any wing having different nonlinear aerodynamic characteristics of airfoil. The present method will contribute to the analysis of aircraft in the conceptual design because the present method can predict the nonlinear aerodynamic characteristics of a wing with a few computing resources and significant time. Copyright © 2007 John Wiley & Sons, Ltd.     

An inviscid/viscous coupling approach for vortex flow field calculations

A new computational approach is developed for the analysis of vortex-dominated flow fields around highly swept wings at high angles of attack. In this approach an inviscid Euler technology is coupled with viscous models, similar to inviscid/boundary layer coupling. The viscous nature of the vortex core is represented by an algebraic model derived from the Navier–Stokes equations. The approach also accounts for the effects of the viscous shear layer near a wing surface through a modified surface boundary condition. The inviscid/viscous coupling consistently provides improved predictions of leading edge separation, vortex bursting and secondary vortex formation at relatively low computational cost. Results for several cases are compared with wind tunnel tests and other Euler and Navier-Stokes solutions.     

飞行器的升力、阻力及升力与环量定理的起源

本文介绍了空气动力学中几个基本概念与定律的起源。其中,升力与阻力分别是空气对物体作用力的两个方向上的分量,它们均是由空气与物体的相对运动而产生的,并与该运动速度的平方成正比。库塔儒可夫斯基升力环量定理给出了翼型升力与翼型绕流之间的关系,开启了20世纪早期各国对翼型性能的研究。同时,鉴于理想流体圆柱绕流无阻力的理论结果与实验观察存在的矛盾开始激发人们对黏性流体运动的研究兴趣,并由此诞生了纳维斯托克斯方程组。而后普朗特提出边界层概念,巧妙解决了局部流动与整体流动的关系问题。针对大展弦比直机翼,普朗特又提出了基于升力线假设的升力线模型,并根据翼型气动数据得到三维机翼的气动性能。     

DERIVATION OF LIFT,DRAG AND THE KUTTA--JOUKOWSKI THEOREM FOR AIRCRAFTS 1)

本文介绍了空气动力学中几个基本概念与定律的起源。其中,升力与阻力分别是空气对物体作用力的两个方向上的分量,它们均是由空气与物体的相对运动而产生的,并与该运动速度的平方成正比。库塔儒可夫斯基升力环量定理给出了翼型升力与翼型绕流之间的关系,开启了20世纪早期各国对翼型性能的研究。同时,鉴于理想流体圆柱绕流无阻力的理论结果与实验观察存在的矛盾开始激发人们对黏性流体运动的研究兴趣,并由此诞生了纳维斯托克斯方程组。而后普朗特提出边界层概念,巧妙解决了局部流动与整体流动的关系问题。针对大展弦比直机翼,普朗特又提出了基于升力线假设的升力线模型,并根据翼型气动数据得到三维机翼的气动性能。     

A numerical investigation into the effects of Reynolds number on the flow mechanism induced by a tubercled leading edge

Leading-edge modifications based on designs inspired by the protrusions on the pectoral flippers of the humpback whale (tubercles) have been the subject of research for the past decade primarily due to their flow control potential in ameliorating stall characteristics. Previous studies have demonstrated that, in the transitional flow regime, full-span wings with tubercled leading edges outperform unmodified wings at high attack angles. The flow mechanism associated with such enhanced loading traits is, however, still being investigated. Also, the performance of full-span tubercled wings in the turbulent regime is largely unexplored. The present study aims to investigate Reynolds number effects on the flow mechanism induced by a full-span tubercled wing with the NACA-0021 cross-sectional profile in the transitional and near-turbulent regimes using computational fluid dynamics. The analysis of the flow field suggests that, with the exception of a few different flow features, the same underlying flow mechanism, involving the presence of transverse and streamwise vorticity, is at play in both cases. With regard to lift-generation characteristics, the numerical simulation results indicate that in contrast to the transitional flow regime, where the unmodified NACA-0021 undergoes a sudden loss of lift, in the turbulent regime, the baseline foil experiences gradual stall and produces more lift than the tubercled foil. This observation highlights the importance of considerations regarding the Reynolds number effects and the stall characteristics of the baseline foil, in the industrial applications of tubercled lifting bodies.     

Steady/unsteady aerodynamic analysis of wings at subsonic,sonic and supersonic Mach numbers using a 3D panel method

This paper treats the kernel function of an integral equation that relates a known or prescribed upwash distribution to an unknown lift distribution for a finite wing. The pressure kernel functions of the singular integral equation are summarized for all speed range in the Laplace transform domain. The sonic kernel function has been reduced to a form, which can be conveniently evaluated as a finite limit from both the subsonic and supersonic sides when the Mach number tends to one. Several examples are solved including rectangular wings, swept wings, a supersonic transport wing and a harmonically oscillating wing. Present results are given with other numerical data, showing continuous results through the unit Mach number. Computed results are in good agreement with other numerical results. Copyright © 2003 John Wiley & Sons, Ltd.     

A 3D mesh deformation technique for irregular in‐flight ice accretion

Aircraft holding around busy airports may be requested to sustain as much as 45 min of icing before landing or being diverted to another airport. In this paper, a three‐dimensional mesh deformation scheme, based on a structural frame analogy, is proposed for the numerical simulation of ice accretion during extended exposure to adverse weather conditions. The goal is to provide an approach that is robust and efficient enough to delay or altogether avoid re‐meshing while preserving (enforcing) nearly orthogonal elements at the highly distorted ice surface. Robustness is achieved by suitably modifying the axial and torsional stiffness components of the frame elements in order to handle large and irregular grid displacements typical of in‐flight icing. Computational efficiency is obtained by applying the mesh displacement to an automatically selected small subset of the entire computational domain. The methodology is validated first in the case of deformations typical of fluid‐structure interaction problems, including wing bending, a helicopter rotor in forward flight, and the twisting of a high‐lift wing configuration. The approach is then assessed for aero‐icing on two swept wings and compared against experimental measurements where available. Copyright © 2015 John Wiley & Sons, Ltd.     

Numerical simulation of flapping‐wing insect hovering flight at unsteady flow

A computational fluid dynamics (CFD) analysis was conducted to study the unsteady aerodynamics of a virtual flying bumblebee during hovering flight. The integrated geometry of bumblebee was established to define the shape of a three‐dimensional virtual bumblebee model with beating its wings, accurately mimicking the three‐dimensional movements of wings during hovering flight. The kinematics data of wings documented from the measurement to the bumblebee in normal hovering flight aided by the high‐speed video. The Navier–Stokes equations are solved numerically. The solution provides the flow and pressure fields, from which the aerodynamic forces and vorticity wake structure are obtained. Insights into the unsteady aerodynamic force generation process are gained from the force and flow‐structure information. The CFD analysis has established an overall understanding of the viscous and unsteady flow around the virtual flying bumblebee and of the time course of instantaneous force production, which reveals that hovering flight is dominated by the unsteady aerodynamics of both the instantaneous dynamics and also the past history of the wing. A coherent leading‐edge vortex with axial flow and the attached wingtip vortex and trailing edge vortex were detected. The leading edge vortex, wing tip vortex and trailing edge vortex, which caused by the pressure difference between the upper and the lower surface of wings. The axial flow, which include the spanwise flow and chordwise flow, is derived from the spanwise pressure gradient and chordwise pressure gradient, will stabilize the vortex and gives it a characteristic spiral conical shape. Copyright © 2006 John Wiley & Sons, Ltd.     

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.     

高升阻比乘波构型优化设计

在M∞ =6, 30km高空条件下,以升阻比为目标函数,进行了锥形流乘波体的黏性优化设计,讨论 了影响乘波体升阻比的因素,并对优化结果进行了数值验证. 结果表明：对于升阻比最大的 黏性优化乘波体,存在最优圆锥角使得源自该基本流场的乘波体升阻比最大;摩阻和波阻处 于同一量级;体积率、细长比和展长比都随着基本流场圆锥角的增大而增大.     

Evaluation of RBFs for volume data interpolation in CFD

A greedy method for choosing an optimum reduced set of control points is integrated with RBF interpolation and evaluated for the purpose of interpolating large‐volume data sets in CFD. Given a function defined at a set of points, the greedy method selects a small subset of these points that is sufficient to keep the interpolation error at all the remaining points below a chosen bound. This is equivalent to a type of data compression and would have useful storage, post‐processing, and computational applications in CFD. To test the method in terms of both the point selection scheme and the suitability of reduced control point volume interpolation, a trial application of the interpolation to velocity fields in CFD volume meshes is considered. To optimise the point selection process, and attempt to be able to capture multiple length scales, a variable support radius formulation has also been included. Structured and unstructured mesh cases are considered for aerofoils, a wing case and a wing‐body case. For smooth volume functions, the method is shown to work well, producing accurate velocity interpolations using a very small number of the cells in the mesh. For general complex fields including large gradients, the method is still shown to be effective, although large gradients require more interpolation points to be used.Copyright © 2013 John Wiley & Sons, Ltd.     

The aerodynamic forces and pressure distribution of a revolving pigeon wing

The aerodynamic forces acting on a revolving dried pigeon wing and a flat card replica were measured with a propeller rig, effectively simulating a wing in continual downstroke. Two methods were adopted: direct measurement of the reaction vertical force and torque via a forceplate, and a map of the pressures along and across the wing measured with differential pressure sensors. Wings were tested at Reynolds numbers up to 108,000, typical for slow-flying pigeons, and considerably above previous similar measurements applied to insect and hummingbird wing and wing models. The pigeon wing out-performed the flat card replica, reaching lift coefficients of 1.64 compared with 1.44. Both real and model wings achieved much higher maximum lift coefficients, and at much higher geometric angles of attack (43°), than would be expected from wings tested in a windtunnel simulating translating flight. It therefore appears that some high-lift mechanisms, possibly analogous to those of slow-flying insects, may be available for birds flapping with wings at high angles of attack. The net magnitude and orientation of aerodynamic forces acting on a revolving pigeon wing can be determined from the differential pressure maps with a moderate degree of precision. With increasing angle of attack, variability in the pressure signals suddenly increases at an angle of attack between 33° and 38°, close to the angle of highest vertical force coefficient or lift coefficient; stall appears to be delayed compared with measurements from wings in windtunnels.     

Characterization of the Mechanics of Compliant Wing Designs for Flapping-Wing Miniature Air Vehicles

Flapping-wing miniature air vehicles (MAVs) offer multiple performance benefits relative to fixed-wing and rotary-wing MAVs. This investigation focused on the problem of designing compliant wings for a flapping-wing MAV where only the spar configuration was varied to achieve improved performance. Because the computational tools needed for identifying the optimal spar configuration for highly compliant wing designs have yet to be developed, a new experimental methodology was developed to explore the effects of spar configuration on the wing performance. This technique optically characterized the wing deformations associated with a given spar configuration and used a customized test stand for measuring lift and thrust loads on the wings during flapping. This revealed that spar configurations achieving large and stable deformed volume during the flapping cycle provided the best combination of lift and thrust. The approach also included a sensitivity and reproducibility analysis on potential spar configurations. Results indicated that the wing shape and corresponding lift and thrust performance were very sensitive to slight changes in volume and 3-D shape associated with slight variations in the spar locations.     

Aerodynamic Performance of a Gliding Swallowtail Butterfly Wing Model

In the present study, we perform a wind-tunnel experiment to investigate the aerodynamic performance of a gliding swallowtail-butterfly wing model having a low aspect ratio. The drag, lift and pitching moment are directly measured using a 6-axis force/torque sensor. The lift coefficient increases rapidly at attack angles less than 10° and then slowly at larger attack angles. The lift coefficient does not fall off rapidly even at quite high angles of attack, showing the characteristics of low-aspect-ratio wings. On the other hand, the drag coefficient increases more rapidly at higher angles of attack due to the increase in the effective area responsible for the drag. The maximum lift-to-drag ratio of the present modeled swallowtail butterfly wing is larger than those of wings of fruitfly and bumblebee, and even comparable to those of wings of birds such as the petrel and starling. From the measurement of pitching moment, we show that the modeled swallowtail butterfly wing has a longitudinal static stability. Flow visualization shows that the flow separated from the leading edge reattaches on the wing surface at α < 15°, forming a small separation bubble, and full separation occurs at α ≥ 15°. On the other hand, strong wing-tip vortices are observed in the wake at α ≥ 5° and they are an important source of the lift as well as the main reason for broad stall. Finally, in the absence of long hind-wing tails, the lift and longitudinal static stability are reduced, indicating that the hind-wing tails play an important role in enhancing the aerodynamic performance.     

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.      

蜻蜓滑翔时柔性褶皱前翅气动特性分析

蜻蜓是自然界优秀的飞行家,滑翔是其常见且有效的飞行模式.蜻蜓优异的飞行能力来源于其翅膀的巧妙结构,褶皱是蜻蜓翅膀上最为显著的结构之一,不仅提高了翅膀的刚度,还改变了其气动特性,而飞行过程中柔性翅膀会产生变形是蜻蜓翅膀的另一特性.为揭示蜻蜓在滑翔时,柔性褶皱前翅的变形,探究褶皱和柔性的共同作用对其气动特性的影响,基于逆向工程,依据前人的测量数据和研究成果,通过三维建模软件建立了蜻蜓三维褶皱前翅的计算流体力学(computational fluiddynamics,CFD)模型和计算结构力学(computational structuralmechanics,CSD)模型,并通过模态分析验证了此模型有足够的精度.基于CFD方法和CFD/CSD双向流固耦合计算方法分别对蜻蜓滑翔飞行时刚性和柔性褶皱前翅的气动特性进行了数值模拟,结果表明,柔性褶皱前翅受气动载荷后,翅脉和翅膜产生形变,柔性前翅上下表面压力差相较于刚性前翅减小了,从而其升力和阻力也减小了,而在大攻角时,变形后的前缘脉诱导出比刚性前翅更强的前缘涡.因此在攻角小于10$^\circ$时刚性前翅的气动特性优于柔性前翅,继续增大攻角,柔性前翅的气动特性则优于刚性前翅.前翅受载后气动响应时间短,翅尖的变形最大,仅仅产生了垂直于翅膀所在平面方向上的变形,而没有发生扭转,翼根处受到应力最大,褶皱上凸部分承受蜻蜓滑翔时前翅的主要载荷.      

跨音速后掠翼抖振边界计算

本文提出一种确定跨音速后掠翼抖振边界的数值计算方法,现有的确定跨音速翼型抖振边界的F.Thomas 准则被推广到包括具有大后掠角的后掠翼,计算是对侧滑翼进行,其中用积分法对三维可压缩湍流边界层的计算是根据本文作者听发展的方法,对于跨音速压强分布是利用A.Eberle的解全速位方程的有限元素法给出,按本文方法计算出F-86A 飞机的抖振边界与相同雷诺数下飞行试验所得结果符合得很好。     

昆虫飞行的空气动力学

昆虫是最早出现、数量最多和体积最小的飞行者. 它们能悬停、跃升、急停、快速加速和转弯, 飞行技巧十分高超. 由于尺寸小, 因而翅膀的相对速度很小, 从而进行上述飞行所需的升力系数很大. 但昆虫翅膀的雷诺数又很低. 它们是如何在低雷诺数下产生高升力的, 是流体力学和生物学工作者都十分关心的问题. 近年来这一领域有了许多研究进展. 该文对这些进展进行综述, 并对今后工作提一些建议. 因2005 年前的工作已在几篇综述文章有了详细介绍, 该文主要介绍2005 年以来的工作. 首先简述昆虫翅的拍动运动及昆虫绕流的基本方程和相似参数; 然后对2005 年之前的工作做一简要回顾. 之后介绍2005 年后的进展, 依次为: 运动学观测; 前缘涡; 翅膀柔性变形及皱褶的影响; 拍动翅的尾涡结构; 翼/身、左右翅气动干扰及地面效应; 微小昆虫; 蝴蝶与蜻蜓; 机动飞行. 最后为对今后工作的建议.      

Navier–Stokes computation of transonic vortices over a round leading edge delta wing

A 3D Navier–Stokes solver has been developed to simulate laminar compressible flow over quadrilateral wings. The finite volume technique is employed for spatial discretization with a novel variant for the viscous fluxes. An explicit three-stage Runge–Kutta scheme is used for time integration, taking local time steps according to the linear stability condition derived for application to the Navier–Stokes equations. The code is applied to compute primary and secondary separation vortices at transonic speeds over a 65° swept delta wing with round leading edges and cropped tips. The results are compared with experimental data and Euler solutions, and Reynolds number effects are investigated.     

An investigation on the lift force of a wing pitching in dynamic stall for a comfort control vessel

On the basis of a comfort control system for ocean vessels, the control forces and moments in the form of lift forces from active wings are of important interest. In an ocean vessel comfort control system, active wings or fins are commonly used and constantly adjust their angles of attack to produce optimal sea-keeping conditions. The unsteady nature of the flow field around a wing, and the behaviour of the generated lift force must be understood in order to optimize the comfort control system. This paper presents experimental data on the flow past a pitching wing, paying particular attention to the lagging effects between the fluid dynamic lift force and the motion of the wing at large angles of attack as a function of peak angle of attack and reduced frequency of oscillation. The range of motion investigated has been chosen according to the applicability of a comfort control wing surface. Numerical data is also included to aid explanation on some of the witnessed phenomena.