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1.
A flow control technique by local vibration is proposed to improve the aerodynamic performance of a typical airfoil NACA 0012. Both wind-tunnel experiments and a large eddy simulation(LES) are carried out to study the effects of local vibration on drag reduction over a wide range of angles of attack. The application parameters of local vibration on the upper surface of the airfoil are first evaluated by numerical simulations.The mounted position is chosen at 0.065–0.09 of chord length from the leading edge.The influence of oscillation frequency is investigated both by numerical simulations and experiments. The optimal frequencies are near the dominant frequencies of shear layer vortices and wake vortices. The patterns of shear vortices caused by local vibration are also studied to determine the drag reduction mechanism of this flow control method. The results indicate that local vibration can improve the aerodynamic performance of the airfoil. In particular, it can reduce the drag by changing the vortex generation patterns.  相似文献   

2.
Strong asymmetrical vortices appear on the leeward of slender body at high angles of attack,which has very unfavorable effect on the stability and control of the aircraft.A method is developed to control the side force of slender body at high angles of attack,and is verified in wind tunnel.A thin-film triangular self-excited oscillation flag is fixed at the tip of the slender body model whose semi-apex angle is 10°.Side force is approximately linearly proportional to roll-setting angle of self-excited oscillation flag at high angles of attack,and the slop of fitting straight line obtained by the least square method is-0.158.The linear relationship between side force and roll-setting angle provides convenience for developing side force control law of slender body at high angles of attack.Experimental data shows that the side force coefficients vary linearly with roll-setting angles when a specific plastic self-excited oscillation flag is used as the control flag.The range of side force coefficient and roll-setting angle are,respectively,-3.2to 3.0 and-20° to 20°.The device is simple,effective,and is of great potential in engineering application.  相似文献   

3.
The massively separated flows over a realistic aircraft configuration at 40?, 50?, and 60?angles of attack are studied using the delayed detached eddy simulation(DDES).The calculations are carried out at experimental conditions corresponding to a mean aerodynamic chord-based Reynolds number of 8.93 × 10~5 and Mach number of 0.088. The influence of the grid size is investigated using two grids, 20.0×10~6cells and 31.0 × 10~6 cells. At the selected conditions, the lift,drag, and pitching moment from DDES predictions agree with the experimental data better than that from the Reynoldsaveraged Navier–Stokes. The effect of angle of attack on the flow structure over the general aircraft is also studied, and it is found that the dominated frequency associated with the vortex shedding process decreases with increasing angle of attack.  相似文献   

4.
The generation of a hairpin vortex from near-wall streamwise vortices is studied via the direct numerical simulation(DNS) of the streak transient growth in the minimal channel flow at Re_τ— 400.The streak profile is obtained by conditionally averaging the DNS data of the fully developed turbulent channel flow at the same Reynolds number.The near-wall streamwise vortices are produced by the transient growth of the streak which is initially subjected to the sinuous perturbation of the spanwise velocity.It is shown that the arch head of the hairpin vortex first grows from the downstream end of the stronger streamwise vortex and then connects with the weaker,opposite-signed streamwise vortex in their overlap region,forming a complete individual hairpin structure.The vorticity transport along the vortex lines indicates that the strength increase and the spatial expansion of the arch head are due to the stretching and the turning of the vorticity vector,respectively.The hairpin packets could be further produced from the generated individual hairpin vortex following the parent-offspring process.  相似文献   

5.
The evolution of low-speed streaks in the turbulent boundary layer of the minimum channel flow unit at a low Reynolds number is simulated by the direct numer- ical simulation (DNS) based on the standard Fourier-Chebyshev spectral method. The subharmonic sinuous (SS) mode for two spanwise-aligned low-speed streaks is excited by imposing the initial perturbations. The possibilities and the physical realities of the turbulent sustaining in the minimal channel unit are examined. Based on such a flow field environment, the evolution of the low-speed streaks during a cycle of turbulent sus- taining, including lift-up, oscillation, and breakdown, is investigated. The development of streamwise vortices and the dynamics of vortex structures are examined. The results show that the vortices generated from the same streak are staggered along the streamwise direction, while the vortices induced by different streaks tilt toward the normal direction due to the mutual induction effect. It is the spatial variations of the streamwise vortices that cause the lift-up of the streaks. By resolving the transport dynamics of enstrophy, the strength of the vortices is found to continuously grow in the logarithmic layer through the vortex stretching mechanism during the evolution of streaks. The enhancement of the vortices contributes to the spanwise oscillation and the following breakdown of the low-speed streaks.  相似文献   

6.
A new idea of drag reduction and thermal protection for hypersonic vehicles is proposed based on the combination of a physical spike and lateral jets for shockreconstruction. The spike recasts the bow shock in front of a blunt body into a conical shock, and the lateral jets work to protect the spike tip from overheating and to push the conical shock away from the blunt body when a pitching angle exists during flight. Experiments are conducted in a hypersonic wind tunnel at a nominal Mach number of 6. It is demonstrated that the shock/shock interaction on the blunt body is avoided due to injection and the peak pressure at the reattachment point is reduced by 70% under a 4° attack angle.  相似文献   

7.
Topological structure and stability of a slender cross flow is discussed by the stability theory of dynamic system. The inner boundary of flow field was limiting streamline and it was proved that the topological structure connected saddles by limiting streamline is stable. It is proved that the development of slender vortices leads to the change of topological structure about cross flow. And it is the change from stable and symmetrical vortices flow pattern to unstable and symmetrical vortices flow pattern, and then to stable and asymmetrical vortices flow pattern due to little disturbance which leads to the development of asymmetrical slender vortices. The influence of disturbance to flowfield structure was discussed by unfolding theory too.  相似文献   

8.
In this paper, Lorentz forces are proved to be able to suppress separation in flows over hydrofoils. Furthermore, a differential equation of pressure distributions on the hydrofoil surface is derived, from which it is found that BVF (boundary vortex flux) σ is a suitable criterion for describing the lift coefficient variations during the electromagnetic control process. According to our numerical results, the periodic variations of lift for a hydrofoil at an attack angle of 17 ° are analyzed and its inherent mechanism is discussed in detail with the concept of BVE On the other hand, the effects of Lorentz force on the hydrofoil's lift are investigated both experimentally and numerically for different magnitudes and locations.  相似文献   

9.
Flexible wings of insects and bio-inspired micro air vehicles generally deform remarkably during flapping flight owing to aerodynamic and inertial forces,which is of highly nonlinear fluid-structure interaction(FSI)problems.To elucidate the novel mechanisms associated with flexible wing aerodynamics in the low Reynolds number regime,we have built up a FSI model of a hawkmoth wing undergoing revolving and made an investigation on the effects of flexible wing deformation on aerodynamic performance of the revolving wing model.To take into account the characteristics of flapping wing kinematics we designed a kinematic model for the revolving wing in two-fold:acceleration and steady rotation,which are based on hovering wing kinematics of hawkmoth,Manduca sexta.Our results show that both aerodynamic and inertial forces demonstrate a pronounced increase during acceleration phase,which results in a significant wing deformation.While the aerodynamic force turns to reduce after the wing acceleration terminates due to the burst and detachment of leading-edge vortices(LEVs),the dynamic wing deformation seem to delay the burst of LEVs and hence to augment the aerodynamic force during and even after the acceleration.During the phase of steady rotation,the flexible wing model generates more ver-tical force at higher angles of attack(40°–60°)but less horizontal force than those of a rigid wing model.This is because the wing twist in spanwise owing to aerodynamic forces results in a reduction in the effective angle of attack at wing tip,which leads to enhancing the aerodynamics performance by increasing the vertical force while reducing the horizontal force.Moreover,our results point out the importance of the fluid-structure interaction in evaluating flexible wing aerodynamics:the wing deformation does play a significant role in enhancing the aerodynamic performances but works differently during acceleration and steady rotation,which is mainly induced by inertial force in acceleration but by aerodynamic forces  相似文献   

10.
The asymmetric vortices over blunt-nose slender body at high angles of attack result in random side force. In this paper, a nose micro-blowing technology is used to control the asymmetric flow. Pressure measurement and particle image velocimetry(PIV) experiments are conducted in a low-speed wind tunnel to research effects of jet flow rate on asymmetric vortices over blunt-nose slender body. The angle of attack of the model is fixed at 50?and the Reynolds number for the experiments is 1.6×10~5 based on diameter of aftbody. A blow hole(5 mm in diameter) on the nose is processed at circumferential angle θ_b= 90?and meridian angle γ_b= 20?with jet momentum ratio C_μ ranging from 5.30×10~(-7) to1.19×10~(-4). Tests are made under two kinds of perturbations. One is called single perturbation with only blow hole and the other is called combined perturbation consists of blow hole and additional granules set on nose. The results show that whether the model has the single perturbation or the combined one,the sectional side force of x/D = 3 varies in the same direction with the increasement of C_μ and remains stable when C_μ is greater than 3.29×10~(-6). But the stable force values are different according to various perturbations. The fact proves that the size and direction of the side force of blunt-nose slender body can be controlled by the nose micro-blowing.  相似文献   

11.
采用数值计算方法对亚音速三角翼纵向及带有小侧滑情况下的流场结构和气动力特性进行了计算。文中给出了三角翼大迎角纵向情况下气动力、机翼前缘分离涡轴线位置和旋涡破裂位置随迎角的变化规律,以及带有横侧小扰动和小侧滑情况下流场结构的非对称性对气动力的影响。计算结果表明与实验结果符合较好。  相似文献   

12.
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.  相似文献   

13.
The numerical investigation has been performed to explore the feasibility of vortex control by leading edge sucking excitation on a delta wing. The results reveal that the flow on the upper surface of the delta wing changes significantly in a wide range of the angle of attack. For the vortical flow at moderate angle of attack, the secondary and tertiary vortices are weakened or suppressed, and the total lift is almost unchanged. For the stalled flow at high angle of attack, the leading edge concentrated vortex is recovered, and the lift is enhanced with increasing suction rate. For the bluff-body flow at even high angles of attack, the lift can still be improved. The concentrated vortex disappears on the upper surface, and the load increment is nearly unchanged along the chordwise direction. The project supported by the National Natural Science Foundation of China (19802018).  相似文献   

14.
Seven hole probe measurement of leading edge vortex flows   总被引:1,自引:0,他引:1  
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.  相似文献   

15.
基于雨燕翅膀的仿生三角翼气动特性计算研究   总被引:1,自引:1,他引:0  
张庆  叶正寅 《力学学报》2021,53(2):373-385
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.   相似文献   

16.
常思源  肖尧  李广利  田中伟  崔凯 《力学学报》2022,54(10):2760-2772
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.   相似文献   

17.
A detailed investigation of the velocity and vorticity fields of a pair of vortices growing over a 75°-sweep delta wing is carried out through LDV measurements of three components of velocity and vorticity. Data are obtained along one of the vortices. The wing is undergoing a ramp-like pitch-up motion. The evolution of the flow field in four planes normal to the free-stream velocity is captured at 100 time instants through the wing motion. The delta wing is pitched through angles of attack ranging from 28° to 68°. From the velocity data at each incidence, the corresponding vorticity field is calculated. Hysteresis effects on vortex development and breakdown are studied through axial velocity and vorticity contours. The topologies of streamlines and vortex lines are compared with the corresponding topologies of the steady case. It is found that vortex breakdown can be detected first by a drastic reduction of the axial velocity. This phenomenon is developing in a non-axisymmetric fashion, beginning at the inboard side of the vortex. This is followed by a reduction of the axial vorticity component and finally by a reversal of the azimuthal vorticity component.This work was supported by the Air Force Office of Scientific Research, Project No. AFOSR-91-0310 and was monitored by Major Daniel Fant.  相似文献   

18.
李立 《力学与实践》2017,39(1):18-24
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.  相似文献   

19.
Experiments on the unsteady nature of vortex breakdown over delta wings   总被引:2,自引:0,他引:2  
 Vortex breakdown location over delta wings is not steady and exhibits fluctuations along the axis of the vortices. Experiments on the nature and source of these fluctuations were carried out. Spectral analysis and other statistical concepts were used to quantify the unsteady behaviour of vortex breakdown location obtained from flow visualization. The fluctuations consist of quasi-periodic oscillations and high-frequency low amplitude displacements. The quasi-periodic oscillations are due to an interaction between the vortices, which cause the antisymmetric motion of breakdown locations for left and right vortices. The oscillations are larger and more coherent as the time-averaged breakdown locations get closer to each other as angle of attack or sweep angle is varied. The frequency of this organized motion is much smaller than the frequency of any other known instabilities. On the other hand, the most probable frequency for the high-frequency small-amplitude fluctuations of breakdown location is in the same range as the frequency of Kelvin–Helmholtz instability of the separated shear layer. A mechanism for the interaction between the vortices causing the oscillations of breakdown location was proposed. When a splitter plate was placed in the symmetry plane of the wing, the large amplitude quasi-periodic oscillations of breakdown location were suppressed. Received: 10 March 1998 / Accepted: 27 October 1998  相似文献   

20.
An effective means of controlling wing leading-edge stall at high angles of attack is deflection of the nose in order to assure shock-free entrance of the stream. A numerical method of computing the angles of nose deflection and the aerodynamic characteristics of a thin wing of arbitrary planform for a shock-free entrance of the steady ideal incompressible fluid stream is elucidated in this paper on the basis of nonlinear wing theory [1]. The problem is solved by the method of discrete vortices. In the computations, the wing and its wake, replaced by a vortex sheet, are modeled by a system of discrete vortices which are nonlinear segments with constant circulation along the length. The angles of deflection of the nose and the aerodynamic characteristics of the wing, including shunting of the free vortices shed from the side and trailing edges, are determined during the computation. Examples of an electronic digital computer are presented.  相似文献   

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