合成射流对钝尾缘翼型气动特性的影响

研究吸力面存在合成射流的情况下,钝尾缘翼型TR-4000-2000流场结构的变化及其升阻力系数等气动特性参数的变化趋势。在相同射流入口速度条件下,采用计算流体力学软件Fluent对相同来流速度不同攻角情况下翼型流场进行非定常数值模拟计算,分析射流前后翼型升阻力系数变化及翼型表面压力的波动状况;在此基础上,对不同射流频率和不同射流速度情况下翼型流场进行模拟计算,寻求最佳射流参数。结果表明,由于射流及尾缘涡的相互作用导致翼型的升阻力特性不断变化,钝尾缘翼型吸力面合成射流有明显的增升减阻效果,在15°攻角时尤为明显,升力系数提高约40%,阻力系数减小约25%。在量纲一射流速度和量纲一射流频率均为1时,射流对翼型的增升减阻效果最佳。     

Investigation on aerodynamics and active flow control of a vertical axis wind turbine with flapped airfoil

A 2D unsteady numerical simulation with dynamic and sliding meshing techniques was conducted to solve the flow around a threeblade Vertical axis wind turbine (VAWT). The circular wakes, strip-like wakes and the shedding vortex structures interact with each other result in an extremely unstable performance. An airfoil with a trailing edge flap, based on the NACA0012 airfoil, has been designed for VAWT to improve flow field around the turbine. Strategy of flap control is applied to regulate the flap angle. The results show that the flapped airfoil has an positive effect on damping trailing edge wake separation, deferring dynamic stall and reducing the oscillating amplitude. The circular wake vortices change into strip vortices during the pitch-up interval of the airfoils. Examination of the flow details around the rotating airfoil indicates that flap control improves the dynamic stall by diminishing the trend of flow separation. Airfoil stall separation has been suppressed since the range of nominal angle of attack is narrowed down by an oscillating flap. Vortices with large intensity over rotational region are reduced by 90 %. The lift coefficient hysteresis loop of flapped airfoil acts as an O type, which represents a more stable unsteady performance. With flap control, the peak of power coefficient has increased by 10 % relative to the full blade VAWT. Obviously, the proposed flapped airfoil design combined with the active flow control significantly has shown the potential to eliminate dynamic stall and improve the aerodynamic performance and operation stability of VAWT.     

合成射流控制参数对大攻角下翼型气动性能的影响

通过数值模拟的方法,对合成射流控制NACA 0012大攻角下翼型流动分离的参数进行了研究.结果表明:对于射流出口宽度为翼型弦长的0.5％,翼型在18°～24°攻角下的流场,当合成射流作用在翼型头部1％弦长位置,吹气速度比为1,无量纲激励频率在1 附近时,可以达到较好的改善翼型整体气动性能的效果.通过对翼型表面压强系数分...     

波浪型结节改形风机翼型的气动性能研究

采用大涡模拟湍流模型对前后缘波浪型结节改形风机翼型在雷偌数5×104下不同攻角的流动控制机理进行了数值研究。研究表明:相比于标准直翼型NACA0012,改形风机翼型在失速区得到了更平缓的升力曲线。在小攻角(α<12°)工况下,改形翼型的升力系数稍小,然而当攻角(α>12°)时,其升力系数明显提高,最高可达37%。改形翼型由于其前后缘沿展向呈正弦波浪型变化,在不同截面处的呈现出明显不同的尾迹结构,从而导致其表面自由剪切层发生扭曲。这种三维涡在其产生、发展以及推移过程中的相互作用,使得其三维尾迹涡结构在失速区能得到很好的控制,从而达到延迟流动分离及减小失速影响的目的。深入研究前后缘波浪型结节改形风机翼型尾迹结构的流动分布及物理特性等,对于揭示前后缘结节改形风机翼型流动控制机理具有非常重要的意义。     

Flow structures around a butterfly-shaped low-aspect-ratio wing

In the present study, we numerically investigate three-dimensional flow structures around a butterfly-shaped low-aspect-ratio wing and their effect on the aerodynamic force at the Reynolds number of 1000 based on the wing chord length and free-stream velocity. When the angle of attack is less than 10°, the flow is steady and fully attached to the upper-wing surface, by which the lift force increases almost linearly with the angle of attack. As the angle of attack further increases, the flow around the wing becomes unsteady and contains the leading-edge, trailing-edge, wing-tip, and hairpin vortices. At these angles of attack, the drag force increases rapidly with increasing angle of attack due to massive separation at the leading edge, but the lift force increases gradually without abrupt fall-off. This is because the wing-tip vortices induce a strong downward flow interacting with the flow separated from the leading edge and delay subsequent vortex roll-up in the downstream. The wing-tip vortices themselves also produce low-pressure regions on the upper-wing surface and thus provide an additional lift force. The flows separated from the leading and trailing edges are eventually combined into pairs of hairpin vortices which travel downstream in the wake. The formation of the hairpin vortices above the upper-wing surface also generates lowpressure regions, and they are another source of the lift force.     

Unsteady boundary layer for a pitching airfoil at low Reynolds numbers

An experimental study was conducted in order to investigate unsteady boundary layers for a pitching airfoil. An NACA0012 airfoil sinusoid-pitched at quarter chord was employed, and its mean angle-of-attack and oscillation amplitude were 0° and 6°, respectively. To explore the unsteady boundary layers, smoke-wire visualization and surface-mounted probe measurements were pursued for three different cases, varying with Reynolds numbers (Re_c=2.3×10⁴, 3.3×10⁴, and 4.8×10⁴). A reduced frequency of 0.1 was identically set in all cases. Results show that in the presented Reynolds number range, the separation bubble dependent on both angle-of-attack and Reynolds number was observed, accompanied with unsteady laminar separation after reattachment. The unsteady laminar separation occurred at the saddle point, which was formed by the two vortices, the wall, and the external flow, and it was independent of reverse flow. This result indicates that the unsteady laminar separation occurs during the process of transition after the reattachment of separated boundary layer for an unsteady flow. The reverse flow observed over the trailing edge significantly interacted with the trailing edge vortex that rotates in the streamwise direction. This trailing edge vortex prevents the uppermost of the reverse flow from reaching to the unsteady laminar separation point during the upstroke, and this induces that the boundary layer breakdown does not occur in spite of the occurrence of laminar separation. The discrete vortices are formed by unsteady laminar separation, and its formation is ultimately affected by the Reynolds number. Consequently, it is obvious that the unsteady boundary layers are ultimately sensitive to Reynolds number in a low Reynolds number regime.     

Effect of leading edge boundary layer thickness on dimple flow structure and separation control

This paper concerns the effect of leading edge boundary layer thickness on dimple flow structure and separation control. A flat plate with adverse pressure gradient sufficient for separation was designed. Large eddy simulation (LES) with dynamic Smagorinsky subgrid model was validated. Dimples with R=0.378, 0.994, 1.453 (R is the ratio of leading edge boundary layer thickness to dimple depth) were investigated. The results show that the horseshoe vortex dominates the dimple flow structure. As R increases, the head of the horseshoe vortex rises further away from the wall and moves downstream. Hairpin vortices in the dimple wake are productions of the horseshoe vortex. Both the horseshoe vortex and the hairpin vortices energize the boundary layer by mixing with the free-stream fluid. As R increases, the separation control effectiveness decreases.     

Numerical study on the effects of a synthetic jet actuator on S809 airfoil aerodynamics at different flow regimes and jet flow angles

Due to importance of active flow control methods for future Horizontal axis wind turbines (HAWT), we numerically investigated the effects of Synthetic jet actuators (SJA) on the separation and aerodynamic lift enhancement of a finite span wing with the S809 airfoil. The investigation was performed at a realistic AOA range (0°~25°), which is associated to a wide range of AOA, from an attached flow to a highly separated condition. Also, effects of SJA injection angle on flow characteristics at a highly separated condition were studied. The unsteady simulations were performed using DES turbulence model. The results show that for an attached flow, the SJA triggers an early separation and reduces the lift coefficients. However, for a separated flow, the SJA considerably enhances lift performance. In addition, at small AOAs the lift values fluctuate harmonically over time; however, as the AOA increases and the separation becomes severe, the lift becomes more chaotic. Finally, a study of the present SJA configuration reveals that the SJA can enhance the lift values at a wide range of injected flow angle. However, lift values do not show distinguished pattern for different jet angles at a high AOA.

     

波浪结构对圆柱-翼型湍流干涉噪声的影响研究

为探究波浪翼型的降噪效果,采用大涡模拟（LES）和边界元法（BEM）相结合的混合方法对3种不同波浪翼型进行模拟,并通过试验验证了仿真模型的可行性,进一步分析了3种不同波浪翼型（表面波浪Wavy airfoil-A、前缘波浪Wavy airfoil-B和前缘+表面波浪Wavy airfoil-C）对圆柱-翼型湍流干涉噪声的影响。研究结果表明：3种模型都能在一定程度上降低翼型湍流干涉噪声,其中Wavy airfoil-C模型降噪效果最好,降噪频率范围最广,其垂直流向方向总声压级降噪量可达6.7 dB;Wavy airfoil-C模型不仅能有效地降低翼型表面压力脉动、各截面上的湍流强度、升阻力系数波动、功率谱密度,还能利用其前缘波浪结构有效地减少前缘主声源区域的面积,且能利用其表面波浪结构的导流作用降低翼型后缘的声源振动幅值。     

钝尾缘风力机翼型气动性能计算分析

钝尾缘风力机翼型有较好的结构和气动性能,是目前多被用于大型风力机叶片靠近轮毂区域的选定翼型.但钝尾缘翼型也有缺点,易产生大的脱流涡,这会降低叶片的气动性能.为了更好地研究钝尾缘翼型的性能,以了解其气动性能的降低能否与其结构性能的优化相匹配.采用计算流体动力学(Computational fluid dynamics,CFD)方法,对薄尾缘翼型S809和改进的钝尾缘翼型S809-100的性能进行模拟和对比,结果表明相对于薄尾缘翼型,钝尾缘翼型可以增大断面的最大升力系数和升力曲线斜率,并可以降低翼型污染对翼型升力影响的敏感度.     

翼身连接处几何参数优化匹配的流动机理仿真

飞机由各个部件组装成为一个整体,在各个连接部位彼此气流互相干扰,会导致干扰阻力增加,严重影响飞机大迎角下的气动性能,因此,优化连接部位几何参数匹配对减小干扰阻力及改善飞机大迎角流动具有重要的工程应用价值.以某一特定飞机的机翼-机身的连接部位着陆构型为研究对象,采用数值模拟的方法,从连接部位的前缘、展向中段及后段3个方面...     

On the aerodynamic forces of a plunging airfoil

The generation of aerodynamic forces by a plunging NACA0012 airfoil at a Reynolds number of 20,000 was studied for a range of non-dimensional plunge frequenciesk and amplitudesh using a 2D unsteady compressible Navier-Stokes solver, an unsteady panel method (UPM) and Garrick’s analysis. Calculations using these two methods indicate that the forces collapse reasonably well withkh (or equivalents the Strouhal number), but are only weak functions ofk. In contrast results from the NS code indicate that the forces are dependent on bothk andkh independently, with large variations at low frequencies. The frequency dependence was found to be a result of vortex shedding from the leading edge of the airfoil, and appears to result from two factors. Firstly at high plunge frequenciesk, the leading edge vortex does not have sufficient time to grow, whereas at lowk the vortex can become a sizeable fraction of the airfoil chord before separating. Secondly once the vortex separates, it is convected downstream over the surface of the airfoil. Due to the low pressure in the vortex core, thrust is maintained while the vortex is upstream of the airfoil maximum thickness point (where the airfoil surface is tilted upstream and the vortex low pressure creates an upstream suction force). Once past this point, the airfoil surface is tilted downstream and the vortex contributes to drag rather than thrust. At high plunge frequencies the vortex cannot be convected far downstream before the motion cycle creates another leading edge vortex on the opposite side of the airfoil, so the impact is lessened. At lowk however the vortex travels far downstream over the airfoil surface, causing drag for a larger portion of the flapping cycle, and therefore lower propulsive efficiency. These results have strong implications on how the thrust and the propulsive efficiency can be maximised by controlling the relative amplitudes and phases of combined pitching and plunging motions of an airfoil.     

Numerical simulation of the flow around two fixed circular airfoils positioned in tandem using the LS-STAG method

In recent years the popularity of immersed boundary methods has been increasing in computational hydrodynamics. One of the most effective methods of this class is the LS-STAG method developed in 2010, which allows computations on sufficiently coarse grids. A software package was developed to solve a number of hydrodynamics and hydroelasticity problems by the LS-STAG method. We present the results of the testing the developed software package by simulating a flow around two fixed circular airfoils positioned in tandem at different distances between the airfoils. We simulate the flow modes for which two symmetric vortices, two asymmetrical vortices, and a vortex wake form behind the front airfoil. For each mode, the typical time dependences of the drag force and lift coefficients are presented. The results agree well with the experimental data in the literature and numerical results of other authors.     

CFD investigation on the flatback airfoil effect of 10 MW wind turbine blade

The flatback airfoil effect on the inboard region of a large wind turbine blade was investigated by numerical analysis. Complicated flow phenomena in wind turbine blade with flatback and non-flatback airfoil were captured by Reynolds-averaged Navier–Stokes flow simulation with shear stress transport turbulence model. Although both airfoil blades were designed using blade element momentum theory to produce identical shaft power, results of three-dimensional computational fluid dynamics (CFD) flow analysis indicated that at a specific location of the root area, the flatback airfoil improved the inboard force by approximately 6 % compared with the non-flatback airfoil. We were also able to confirm that by using the flatback airfoil, the overall shaft power throughout the blade increased by 1 %, thereby restraining the bending moment exerted by the thrust force on the hub by 0.5 %. Moreover, numerical analysis results indicated that the flatback airfoil blade reduced the size of the secondary vortex around the blade root area and its progress in the secondary direction in comparison with the non-flatback airfoil blade. The shape of the flatback airfoil on the trailing edge weakened the adverse pressure gradient migrating from the lower to the upper surface. Regardless of the flatback airfoils, the tip vortex core of the outboard region formed on the suction surface leading edge and strongly rolled up by the pressure surface boundary layers due to the large pressure difference between the suction and pressure surfaces in the blade tip region. This remarkable strong tip vortex developed downstream and raked up the boundary layer of the blade trailing edge with low energy.     

Unsteady flow field numerical calculations of a ship’s rotating Weis-Fogh-type propulsion mechanism with the advanced vortex method

The Weis-Fogh mechanism, found in the hovering flight of a small bee, is a unique and efficient lift generation. In this study, we proposed a rotating type propulsion model that applies the principle of the Weis-Fogh mechanism and calculated the unsteady flow field of the propulsion model with the advanced vortex method. The wing (NACA0010 airfoil) and channel are approximated by source and vortex panels, and free vortices are introduced away from the body surfaces. The viscous diffusion of fluid is represented using the core-spreading model to the discrete vortices. We investigated the thrust and drag coefficients, pressure field, vorticity field, velocity vector field, and average propulsive efficiency of the propulsion model by changing the rotating angle velocity. The force acting on the wing depended heavily on the directions of the thrust and drag and the thrust and drag coefficients largely fluctuated with the change in the rotating angles. The average thrust increased as the rotating angle velocity increased. The maximum propulsive efficiency was 27.9% at a calculated angle velocity. The flow field of this rotating type propulsion mechanism is unsteady and very complex because the wing rotates and moves unsteadily in the channel. However, using the advanced vortex method, it could be calculated accurately.     

Aerodynamic performance improvement by divergent trailing edge modification to a supercritical airfoil

A computational study has been performed to determine the effects of divergent trailing edge (DTE) modification to a supercritical airfoil in transonic flow field. For this, the computational result with the original DLBA 186 supercritical airfoil was compared to that of the modified DLBA 283. A Navier-Stokes code, Fluent 5. 1, was used with Spalart-Allmaras’s one-equation turbulence model. Results in this study showed that the reduction in drag due to the DTE modification is associated with weakened shock and delayed shock appearance. The decrease in drag due to the DTE modification is greater than the increase in base drag. The effect of the recirculating flow region on lift increase was also observed. An airfoil with DTE modification achieved the same lift coefficient at a lower angle of attack while giving a lower drag coefficient. Thus, the lift-to-drag ratio increases in transonic flow conditions compared to the original airfoil. The lift coefficient increases considerably whereas the lift slope increases just a little due to DTE modification.     

Vortical flows over a delta wing at high angles of attack

The vortex flow characteristics of a sharp-edged delta wing at high angles of attack were studied using a computational technique. Three dimensional, compressible Reynolds-averaged Navier-Stokes equations were solved to understand the effects of the angle of yaw, angle of attack, and free stream velocity on the development and interaction of vortices and the relationship between suction pressure distributions and vortex flow characteristics. The present computations gave qualitatively reasonable predictions of vortical flows over a delta wing, compared with past wind tunnel measurements. With an increase in the angle of yaw, the symmetry of the pair of leading edge vortices was broken and the vortex strength was decreased on both windward and leeward sides. An increase in the free stream velocity resulted in stronger leading edge vortices with an outboard movement.     

涡流发生器对动态失速影响的模拟研究

基于风力机专用翼型DU91-W2-250直叶片段,采用延迟分离涡模拟（Delayed detached-eddy simulation,DDES）与k-ω SST两种CFD模拟方法,研究了涡流发生器（Vortex generators,VGs）对动态失速的影响,并对两种模拟方法的表现进行了比较。结果表明：动态失速下,光滑叶片段的分离与重新附着均出现延迟;加VGs后,上仰阶段叶片段的气动分离更加延迟,下俯阶段叶片段上表面附着流动重建的更早,气动力的迟滞现象得到明显改善。VGs对DU91-W2-250叶片段增升减阻的效果明显,其中最大升力系数增加10%,周期平均升力系数增加17.9%,最大阻力系数减少56.3%,周期平均阻力系数减少40.3%。DDES模型能更细致地反映VGs对动态失速抑制作用下叶片段表面复杂流动分离现象,k-ω SST模型则难以捕捉小尺度涡结构,模拟得到的涡结构展向呈二维分布。     

Numerical study of blowing and suction slot geometry optimization on NACA 0012 airfoil

The effects of jet width on blowing and suction flow control were evaluated for a NACA 0012 airfoil. RANS equations were employed in conjunction with a Menter’s shear stress turbulent model. Tangential and perpendicular blowing at the trailing edge and perpendicular suction at the leading edge were applied on the airfoil upper surface. The jet widths were varied from 1.5% to 4% of the chord length, and the jet velocity was 0.3 and 0.5 of the free-stream velocity. Results of this study demonstrated that when the blowing jet width increases, the lift-to-drag ratio rises continuously in tangential blowing and decreases quasi-linearly in perpendicular blowing. The jet widths of 3.5% and 4% of the chord length are the most effective amounts for tangential blowing, and smaller jet widths are more effective for perpendicular blowing. The lift-to-drag ratio improves when the suction jet width increases and reaches its maximum value at 2.5% of the chord length.     

Jet flow characteristics of sinusoidal wavy nozzles

The flow characteristics of jets issued from a sinusoidal nozzle with in-phase and 180° out-of-phase exit configurations were investigated using PIV (particle image velocimetry) and flow visualization techniques. The experiments were carried out at a Reynolds number of about 6300 based on the mean width of the jet nozzle. Compared to a normal rectangular jet, the sinusoidal nozzle jets have smaller velocity deficits as the flow goes downstream. In addition, the turbulence intensity is suppressed in the horizontal center plane. For the case of in-phase wavy nozzle jet, the length of the potential core exhibits small variations along the lateral direction, while the 180° out-of-phase wavy nozzle jet shows large lateral variation in the length of potential core. The turbulent kinetic energy of the 180° out-ofphase nozzle jet also shows sinusoidal variation in the horizontal planes. Large-scale vortices shed from the sinusoidal edge of the nozzle interact strongly and migrate toward the center plane as the flow develops downstream.