基于雨燕翅膀的仿生三角翼气动特性计算研究

针对低雷诺数微型飞行器的气动布局, 设计出类似雨燕翅膀的一组具有不同前缘钝度的中等后掠($\varLambda =50^{\circ}$)仿生三角翼. 为了定量对比研究三角翼后缘收缩产生的气动效应, 设计了一组具有同等后掠的普通三角翼. 为了深入研究仿生三角翼布局的前缘涡演化特性以及总体气动特性, 采用数值模拟方法详细地探索了低雷诺数($Re=1.58\times 10^{4})$流动条件下前缘涡涡流结构和气动力随迎角的变化规律. 分析结果表明, 前缘钝度和后缘收缩对仿生三角翼前缘涡的涡流强度和涡破裂位置有显著影响. 相对于钝前缘来说, 尖前缘使仿生三角翼上下表面的压力差增大, 涡流强度也更大, 增升作用也更显著. 相对于普通三角翼构型, 仿生三角翼的前缘斜切使其阻力更大, 但后缘的收缩使涡破裂位置固定在此位置, 因此整个上翼面保持低压, 总的升力更大. 由于小迎角时升力增大更明显, 因此仿生三角翼的气动效率在小迎角时明显大于普通三角翼. 这些结论对于揭示鸟类的飞行机理以及未来微型仿生飞行器的气动布局设计具有重要的研究价值.      

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.

     

65°三角翼亚音速复杂流场计算和数据可视化

提出一种基于非结构混合网格和有限体积法的有效计算策略,对第二期国际涡流试验项目(second international vortex flow experiment,VFE-2)的尖前缘65°三角翼在马赫数0.4,迎角20.3°,雷诺数2×10~6条件下的亚音速复杂流场结构进行数值模拟,重点探讨了基于计算数据进行该类型复杂涡系干扰表面和空间流场关键特征提取和数据可视化问题.通过与相关试验类比,建立了与先进试验流动显示技术相比拟的定性和定量分析方法,为三角翼这类复杂流场结构的精细分析奠定了技术基础.采用上述方法,细致分析了亚音速三角翼的大迎角复杂旋涡流场结构,得到了与试验一致的结论.研究证实:在大迎角条件下,三角翼流动物理复杂,黏性效应耦合严重,只有通过N-S方程计算才能准确地捕捉主涡和二次涡的发展.     

偏转矢量喷流对双三角翼前缘涡破裂影响的实验研究

本文应用染色液流动显示技术对后缘偏转喷流情况下76°/40°双三角翼前缘涡破裂位置的变化进行了观测,实验结果表明偏转喷流主要推迟与喷流方向相同一侧前缘涡的破裂,而使另一侧前缘涡破裂略有提前.随着喷流偏转角度的增大,喷流使两前缘涡破裂位置差逐渐增大.另外,随着模型攻角的增大,前缘涡涡核与双三角翼翼面的夹角逐渐增大,导致偏转喷流的作用逐渐减弱.     

昆虫拍翼方式的非定常流动物理再探讨

基于提出的理论模化方法来探讨昆虫拍翼方式的非定常流动物理. 以悬停飞行为 例，通过对拍翼运动的分析，不仅解释了昆虫利用高频拍翼的方式为何能够克服低雷诺数带 来的气动局限性(St \gg 1/Re)，而且还指出高升力产生和调节的3个流动 控制因素：(1) 由于拍翼的变速运动即时引起了流体动力响应，这种附加惯性效应 可产生瞬时的高升力; (2) 保持前缘涡不脱离翼面有助于减少升力的下降; (3) 增大后缘涡的强度并加速其脱离后缘能够有效地提高升力.     

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.     

Computational modeling of vortex breakdown control on a delta wing

We present an effort to model the development and the control of the vortex breakdown phenomenon on a delta wing. The pair of the vortices formed on the suction side of a delta wing is the major contributor to the lift generation. As the angle of attack increases, these vortices become more robust, having high vorticity values. The critical point of a delta wing operation is the moment when these vortices, after a certain angle of attack, are detached from the wing surface and wing stall occurs. In order to delay or control the vortex breakdown mechanism, various techniques have been developed. In the present work, the technique based on the use of jet-flaps is numerically investigated with computational fluid dynamics by adopting two eddy-viscosity turbulence models. The computational results are compared with the experimental data of Shih and Ding (1996). It is shown that between the two turbulence models, the more advanced one, which adopts a non-linear constitutive expression for the Reynolds-stresses, is capable to capture the vortex breakdown location for a variety of jet exit angles. The performance assessment of the models is followed by the investigation of the effect of the jet-flap on the lift and drag coefficients.     

基于三路舵机驱动的扑翼飞行模型气动特性研究

扑动而形成非定常气动现象是扑翼飞行过程中产生高升力的主要原因。本文以Ellington实验的鹰蛾翅膀为原形,设计扑翼实验及数值计算模型。通过压差传感器对翅膀模型上翼面固定位置进行测压,分析前缘涡的产生及脱落情况(考虑动压效应)。测量上下翼面固定位置处的压差,揭示扑翼飞行中产生高升力的主要原因。利用烟风洞观察扑翼模型周围流场结构及特殊涡产生变化情况。另外,根据Ellington提供的升力关系式估算了扑翼模型在一个周期内的平均升力。最后,基于三维欧拉方程对扑翼飞行气动特性进行数值模拟,计算结果与实验吻合良好。     

基于涡量矩理论的绕振荡水翼涡动力学分析

文章采用标准k-ω SST湍流模型和动网格技术, 实现了绕俯仰振荡NACA66水翼非定常流动结构与水动力特性的数值模拟, 并基于有限域涡量矩理论定量表征了局部旋涡结构对水翼动力特性的影响. 研究结果表明: 在水翼升程阶段, 当攻角较小时, 层流向湍流的转捩点由水翼尾缘向前缘移动; 在较大攻角时, 顺时针尾缘涡?TEV在水翼吸力面上生成并向前缘发展, 同时与吸力面上的顺时针前缘涡?LEV融合发展为附着在整个吸力面上的新前缘涡?LEV, 新的?LEV与逆时针尾缘涡+TEV相互作用直至完全脱落, 直接导致了水翼的动力失速, 在回程阶段, 绕振荡水翼的流场结构逐渐由湍流转变为层流. 基于有限域涡量矩理论的定量分析发现, 有限域内附着的?LEV和?TEV提供正升力, 当?LEV发展覆盖整个吸力面时对升力的贡献最大, 占总升力近50%, 而+TEV提供负升力. 同时发现, 有限域内各旋涡内部的不同区域提供的升力有正有负; 而逸出有限域的旋涡内部不同区域提供的升力方向均保持一致, 其中顺时针涡提供正升力, 而逆时针涡提供负升力. 在失速阶段, 域外旋涡整体对升力贡献较小且存在小幅波动, 体现了流动的非定常性.      

Reynolds number effects on leading edge vortex development on a waving wing

The waving wing experiment is a fully three-dimensional simplification of the flapping wing motion observed in nature. The spanwise velocity gradient and wing starting and stopping acceleration that exist on an insect-like flapping wing are generated by rotational motion of a finite span wing. The flow development around a waving wing at Reynolds number between 10,000 and 60,000 has been studied using flow visualization and high-speed PIV to capture the unsteady velocity field. Lift and drag forces have been measured over a range of angles of attack, and the lift curve shape was similar in all cases. A transient high-lift peak approximately 1.5 times the quasi-steady value occurred in the first chord length of travel, caused by the formation of a strong attached leading edge vortex. This vortex appears to develop and shed more quickly at lower Reynolds numbers. The circulation of the leading edge vortex has been measured and agrees well with force data.     

翼反角对高压捕获翼构型亚声速气动特性影响分析研究

高压捕获翼新型气动布局在高超声速设计状态下具有较好的气动性能, 新升力面的引入使其在亚声速条件下也具有较大的升力, 但在亚声速下的稳定特性还有待研究. 基于高压捕获翼气动布局基本原理, 在机身-三角翼组合体上添加单支撑和捕获翼, 设计了一种参数化高压捕获翼概念构型. 以捕获翼和机体三角翼上/下反角为设计变量, 采用均匀试验设计、计算流体力学数值计算方法及Kriging代理模型方法, 研究了0° ~ 10°攻角状态下不同翼反角对高压捕获翼构型亚声速气动特性的影响, 重点分析了升阻特性、纵向和横航向稳定性的变化规律以及流场涡结构等. 结果表明, 小攻角状态下翼反角对升阻比的影响比大攻角更加显著, 捕获翼上反时, 升阻比略微增大, 下反则升阻比减小; 三角翼上反时, 升阻比减小, 下反则升阻比先略微增大后缓慢减小; 翼反角对纵向稳定性的总体影响较小, 捕获翼上反会稍微提高纵向稳定性, 而三角翼上反则会降低纵向稳定性; 捕获翼或三角翼上反都会增强横向稳定性, 下反则减弱横向稳定性, 但大攻角状态时, 三角翼上反角过大对提升横向稳定性作用有限; 捕获翼上反航向稳定性增强, 下反航向稳定性则减弱, 而三角翼下反对提升航向稳定性的整体效果比上反更加显著.      

A Numerical Study of the Flow Around a Model Winged Seed in Auto-Rotation

In this study the flow around a winged-seed in auto-rotation is characterized using direct numerical simulations (DNS) at Reynolds number in the range 80–240, based on the descent speed and a characteristic chord length. In this range, the flow is approximately steady when observed from a reference frame fixed to the seed. For all cases, the flow structure consists of a wing tip vortex which describes a helical path, a vortex shed behind the nut of the seed and a stable leading edge vortex above the wing surface which merges with the tip vortex. With increasing Reynolds number, the leading edge vortex becomes more intense and gets closer to the wing surface. The simulation results also show the formation of a spanwise flow on the upper surface of the wing, moving fluid towards the wing tip in a region downstream and beneath the leading edge vortex. This spanwise flow is rather weak inside the core of the leading edge vortex, and the analysis of the streamlines show a very weak transport of vorticity along the vortex for the cases under consideration. The analysis of the flow suggests that the stabilization of the leading edge vortex is mainly due to non-inertial accelerations, although viscous effects may contribute, specially at lower Re. Furthermore, the leading edge vortex has been characterized by analysing the flow variables averaged along cross-sections of the vortex. While some quantities, like the spanwise velocity or the pressure inside the vortex, are rather insensitive to the threshold used to define the leading edge vortex, the same is not true for the circulation of the vortex or its averaged spanwise vorticity, due to the viscous nature of the vortex. Finally, it is observed that the spanwise vorticity scales with the angular rotation of the seed for the different Re.     

The flexibility effect of a plate according to various angles of attack in a free-stream

This paper presents a computational fluid–structure interaction analysis for a flexible plate in a free-stream to investigate the effects of flexibility and angle of attack on force generation. A Lattice Boltzmann Method with an immersed boundary technique using a direct forcing scheme model of the fluid is coupled to a finite element model with rectangular bending elements. We investigated the effects of various angles of attack of a flexible plate fixed at one of the end edges in a free-stream at a Reynolds number of 5000, which represents the wing flapping condition of insects and small birds in nature. The lift of the flexible plate is maintained at the large angle of attack, whereas the rigid plate shows the largest lift at angles of attack around 30–40° and then drastic reductions in the lift at the large angle of attack. If we consider the efficiency as the lift divided by the drag, the flexible plate shows better efficiency at angles of attack greater than 30° compared to the rigid plate. The better performance of the flexible plate at large angles of attack comes from the deformation of the plate, which produces an interaction between the trailing edge vortex and the short edge vortex. The horseshoe-shaped vortex produced by a large vortex interaction at the trailing edge side has an important role in increasing the lift, and the small projection area due to the deformation reduces the drag. Furthermore, we investigate the role of flexibility on the lift and the drag force of the rectangular plate in a free-stream as the Reynolds number increases. Whenever a large vortex interaction at the trailing edge side is shown, the efficiency of the rectangular plate is improved. Especially, the flexible plate shows better efficiency as the Reynolds number increases regardless of the angle of attack.     

Seven hole probe measurement of leading edge vortex flows

This paper discusses the use of a seven-hole probe on measurements of leading edge vortices of highly sweep delta wing planforms. Intrusive probe data taken with the pressure probe were compared with non-intrusive measurements made with laser Doppler anemometry system. In addition to probe size, the natural position of breakdown and the sweep angle of the wing are also factors in determining sensitivity of the flow to probe interference. At low angles of attack vortex breakdown does not occur in the vincinity of the model and the seven hole probe was found to yield reasonably accurate measurements. When the angle of attack of the model was increased so that vortex breakdown was near the trailing edge, introducing the probe over the wing would cause the breakdown position to move ahead of the probe. However, when breakdown naturally occurred ahead of the mid-chord of the wing the vortices were found to be less sensitive to a probe placed behind the breakdown point. Vortex breakdown on a lower swept wing is found to be more sensitive to interference. Near the breakdown region, seven hole probe measurement is less accurate due to a combination of probe interference and flow reversal.     

Application of a vortex lattice numerical model in the calculation of inviscid incompressible flow around delta wings

The flow over a flat plate delta wing at incidence and in sideslip is studied using vortex lattice models based on streamwise penelling. For the attached flow problem the effect of sideslip is simulated by modifying the standard vortex lattice model for zero sideslip by aligning the trailing vortices aft of the wing along the resultant flow direction. For the separated flow problem a non-linear vortex lattice model is developed for both zero and non-zero sideslip angles in which the shape and position of the leading edge separation vortices are calculated by an iterative procedure starting from an assumed initial shape. The theoretical values are compared with available theoretical and experimental results.     

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.      

Feedback control of vortex shedding from an inclined flat plate

Open- and closed-loop control of vortex shedding in two-dimensional flow over a flat plate at high angle of attack is numerically investigated at a Reynolds number of 300. Unsteady actuation is modeled as a body force near the leading or trailing edge and is directed either upstream or downstream. For moderate angles of attack, sinusoidal forcing at the natural shedding frequency results in phase locking, with a periodic variation of lift at the same frequency, leading to higher unsteady lift than the natural shedding. However, at sufficiently high angles of attack, a subharmonic of the forcing frequency is also excited and the average lift over the forcing period varies from cycle-to-cycle in a complex manner. It is observed that the periods with the highest averaged lift are associated with particular phase differences between the forcing and the lift, but that this highest-lift shedding cycle is not always stably maintained with open-loop forcing. We design a feedback algorithm to lock the forcing with the phase shift associated with the highest period-averaged lift. It is shown that the compensator results in a stable phase-locked limit cycle for a broader range of forcing frequencies than the open-loop control, and that it is able to stabilize otherwise unstable high-lift limit cycles that cannot be obtained with open-loop control. For example, at an angle of attack of 40°, the feedback controller can increase the averaged magnitude of force on the plate by 76% and increase the averaged lift coefficient from 1.33 to 2.43.     

激光片光流动显示技术在高机动飞机布局设计中的应用

采用激光片光流动显示技术,针对某高机动布局飞机开展大迎角流动机理研究.结果表明激光片光流动显示技术具备捕捉复杂流动结构的能力,可以清晰观察双前翼布局大迎角下机头涡、第一前翼涡、第二前翼涡以及机翼涡等涡结构的生成、发展、融合等复杂流动结构的演化,并且揭示出涡结构扫略垂尾内侧使布局航向更加安定,偏转前翼将使旋涡绕流的涡心降低,造成布局升力提高,低头力矩增加.片光流动显示结果进一步表明,布局航向安定性不足,垂尾需要沿展向外移动,为布局深化设计提供了设计依据.     

Computation of field structure and aerodynamic characteristics of delta wings at high angles of attack

A numerical investigation of the structure of the vortical flowfield over delta wings at high angles of attack in longitudinal and with small sideslip angle is presented. Three-dimensional Navier-Stokes numerical simulations were carried out to predict the complex leeward-side flowfield characteristics that are dominated by the effect of the breakdown of the leading-edge vortices. The methods that analyze the flowfield structure quantitatively were given by using flowfield data from the computational results. In the region before the vortex breakdown, the vortex axes are approximated as being straight line. As the angle of attack increases, the vortex axes are closer to the root chord, and farther away from the wing surface. Along the vortex axes, as the adverse pressure gradients occur, the axial velocity decreases, that is, A is negativee, so the vortex is unstable, and it is possible to breakdown. The occurrence of the breakdown results in the instability of lateral motion for a delta wing, and the lateral moment diverges after a small perturbation occurs at high angles of attack. However, after a critical angle of attack is reached the vortices breakdown completely at the wing apex, and the instability resulting from the vortex breakdown disappears.     

Vortex structure of separated flows on model wings at low freestream velocities

Flow past model wings is experimentally investigated in a subsonic wind tunnel at large angles of attack at which the laminar boundary layer separates near the leading edge of the wing (flow stall). The object of the study was the flow structure within the separation zone. The carbon-oil visualization of surface streamlines used in the experiments showed that in the separation zone there exist one or more pairs of large-scale vortices rotating in the wing plane. Certain general properties of the vortex structures in the separation zone are found to exist, whereas the flow patterns may differ depending on the model aspect ratio, the yaw angle, and other factors.