Planform and Flexibility on Lift Characteristics for Flow over Low-Aspect Ratio Wings

Lift enhancement through flexibility of plunging wings at low Reynolds numbers

Insect wings are subjected to fluid, inertia and gravitational forces during flapping flight. Owing to their limited rigidity, they bent under the influence of these forces. Numerical study by Hamamoto et al. (Adv Robot 21(1–2):1–21, 2007) showed that a flexible wing is able to generate almost as much lift as a rigid wing during flapping. In this paper, we take a closer look at the relationship between wing flexibility (or stiffness) and aerodynamic force generation in flapping hovering flight. The experimental study was conducted in two stages. The first stage consisted of detailed force measurement and flow visualization of a rigid hawkmoth-like wing undergoing hovering hawkmoth flapping motion and simple harmonic flapping motion, with the aim of establishing a benchmark database for the second stage, which involved hawkmoth-like wing of different flexibility performing the same flapping motions. Hawkmoth motion was conducted at Re = 7,254 and reduced frequency of 0.26, while simple harmonic flapping motion at Re = 7,800 and 11,700, and reduced frequency of 0.25. Results show that aerodynamic force generation on the rigid wing is governed primarily by the combined effect of wing acceleration and leading edge vortex generated on the upper surface of the wing, while the remnants of the wake vortices generated from the previous stroke play only a minor role. Our results from the flexible wing study, while generally supportive of the finding by Hamamoto et al. (Adv Robot 21(1–2):1–21, 2007), also reveal the existence of a critical stiffness constant, below which lift coefficient deteriorates significantly. This finding suggests that although using flexible wing in micro air vehicle application may be beneficial in term of lightweight, too much flexibility can lead to deterioration in flapping performance in terms of aerodynamic force generation. The results further show that wings with stiffness constant above the critical value can deliver mean lift coefficient almost the same as a rigid wing when executing hawkmoth motion, but lower than the rigid wing when performing a simple harmonic motion. In all cases studied (7,800 ≤ Re ≤ 11,700), the Reynolds number does not alter the force generation significantly.

Micro air vehicles (MAVs), a class of unmanned air vehicles, usually have low aspect ratio (AR) wings and operate in the range of low Reynolds number (Re). In this study, the effect of AR and Re on the performance of low AR wings have been carried out in a subsonic wind tunnel to generate lift, drag and pitching moment data. This obtains sufficient insights into the understanding of MAV aerodynamics. The force measurements are done for the low AR wings composed of S5010 and E214. The AR in each case was varied in the range of 0.5 to 2.0, whereas the Re was varied in the range of 80 × 10 3 to 120 × 10 3 . The effect of Re on the lift is found to be more prominent in the post-stall region for all ARs except 0.5. This study also indicates that with a decrease in AR, the impact of Re on performance becomes less pronounced. As the AR reduces, the maximum lift coefficient and the lift-to-drag ratio of the wings drop significantly. Similarly, the slope of the lift curve also decreases with a decrease of AR. Furthermore, for a given AR, the induced drag factor of both wings decreases with increasing Re, while at a fixed Re, it increases as AR decreases. In contrast, the stall angle shows an improvement with the reduction of AR and it is more prominent for the wing of E214 airfoil. The pitching moment values become more negative with an increase of AR, indicating the pitch-down tendency of the wing to rise with AR.

SummaryAn experimental investigation has been performed to study the formation and development of spiral vortex flow over a swept-back wing. An aerofoil section with three alternative leading edge shapes was tested at sweep angles ranging from 0° to 56° for unit Reynolds numbers of 1 × 106/m and 2 × 106/m. The principal diagnostic tool was the surface oil-flow visualisation technique supplemented by pressure distribution measurements in certain cases. No spiral vortex flow was observed for sweep angles of 0° and 15° but at higher sweep angles the oil-flows indicated that there were three different mechanisms for the formation of spiral vortices. The angle of incidence at the onset of vortex flow, and the mechanism responsible for its formation, were found to depend upon the sweep angle, the leading edge shape and the Reynolds number. It was also noted that the larger the leading edge radius the greater the dependence upon Reynolds number. However, comparison with other work suggests that Reynolds number, incidence and sweep angle alone are insufficient to determine the type of spiral vortex flow occurring on a given wing.

<p>Flow control has significant technological importance as it can manipulate the flow field in a desired way either actively or passively. This wide research area has remained the point of attention for many years as it is applicable to various applications. Blowing as a flow control method, among other methods, is more technically feasible and market ready technique. A brief review from the existing literature on various studies on blowing has been presented along with their outcome. Then, studies were conducted to investigate the performance variation (in terms of lift coefficient, drag coefficient) of different airfoils with respect to various blowing parameters. It was observed that for NACA 0012 airfoil the maximum lift coefficient peaks at a blowing ratio 0.2 and then it decreases whereas the stalling angle increases with rise of blowing ratio. For LA203A both maximum lift coefficient and stalling angle increases with blowing ratio. For thick airfoil, it was found that lift increases with the rise of moment coefficient and blowing ratio and mid chord slots gives better performance at lower angles of attack whereas leading edge slots exhibits better performance at higher angles of attack. In case of thick elliptical airfoil, increase in lift coefficient was noticed with the increase in moment coefficient and blowing ratio but an optimum jet width to chord ratio (0.41) was found beyond which increase in jet width causes drop in lift coefficient. Additionally, it was found that at lower blowing angle thick elliptical airfoil performs better compared to that of higher angles of attack. Study on NACA 0012 and Aerospatiale A proved the fact that lifts increases with rise in blowing ratio at three different jet diameters to chord ratios. This study also found that higher blowing ratio is also advantageous for the turbo machinery in increasing the pressure difference. Lastly, the study on low Re airfoil flow found that increasing blowing ratio has a deleterious effect on aerodynamics performance (in terms of lift and drag coefficient).</p>

Micro air vehicles (MAVs) with a wingspan of 15cm or shorter, and flight speed around 10m∕s have attracted substantial interest in recent years. There are several prominent features of MAV flight: (i) low Reynolds number (104-105), resulting in degraded aerodynamic performance, (ii) small physical dimensions, resulting in certain favorable scaling characteristics including structural strength, reduced stall speed, and impact tolerance, and (iii) low flight speed, resulting in order one effect of the flight environment and intrinsically unsteady flight characteristics. Flexible wings utilizing membrane materials are employed by natural flyers such as bats and insects. Compared to a rigid wing, a membrane wing can better adapt to the stall and has the potential for morphing to achieve enhanced agility and storage consideration. We will discuss the aerodynamics of both rigid and membrane wings under the MAV flight condition. To understand membrane wing performance, the fluid and structure interaction is of critical importance. Flow structures associated with the low Reynolds number and low aspect ratio wing, such as pressure distribution, separation bubble, and tip vortex, as well as structural dynamics in response to the surrounding flow field are discussed. Based on the computational capabilities for treating moving boundary problems, an automated wing shape optimization technique is also developed. Salient features of the flexible-wing-based MAV, including the vehicle concept, flexible wing design, novel fabrication methods, aerodynamic assessment, and flight data analysis are highlighted.

The aerodynamics of the low aspect ratio (LAR) wings is of outmost importance in the performance of the fixed-wing micro air vehicles (MAVs). The flow around these wings is widely influenced by three dimensional (3D) phenomena: including wing-tip vortices, formation of laminar bubble, flow separation and reattachment, laminar to turbulent transition or any combination of these phenomena. All the recent studies consider the aerodynamic characteristics of the LAR wings under the effect of the direct wind. Here we focus on the numerical study of the influence of cross-wind on flow over the inverse Zimmerman wings with the aspect ratios (AR) between 1 and 2 at Reynolds numbers between 6×104 and 105. We have considered cross-wind’s angles from 0° to 40° and angle of attack from 0° to 12°. The results show that lift and drag coefficient generally decrease when the angle of the cross-wind is increased.

<abstract> Ice accretion can reduce the performance of aircraft's wings, which results in higher fuel consumption and risk of accidents. Experiments proved that even in its very earlier stages (increased roughness), icing could cause a reduction of 25% in the maximum lift, and an increase of 90% in drag of an aspect ratio 6 wing. In this work, we propose a correlation to predict the degradation of the maximum lift coefficient caused by roughness effects on flows over two airfoils, a NACA 0012 and a model 5–6. In addition, a second correlation is proposed to find the minimum Reynolds number that are useful for higher Reynolds number applications when roughness is considered. The SA roughness extension is implemented into an open-source code called SU2. The verification and validation of the implementation is performed in two steps. First, the behavior of the flow over a smooth NACA 0012 is investigated to confirm whether the implementation has no influence on the original model when roughness is not activated. Then, roughness is activated, and estimations of lift coefficients and velocity profiles inside the boundary layer are evaluated and compared to numerical and experimental results. Finally, investigations on the maximum lift coefficient reduction caused by different equivalent sand grain roughness heights and Reynolds numbers are performed. Our results demonstrated that, for the equivalent sand grain roughness heights investigated, the variation of sufficiently small heights has no significant influence on the maximum lift coefficient degradation. Moreover, when roughness is continuously increased, a saturation point seems to be approached, in which the variation of the maximum lift coefficient degradation is reduced. We noticed that although the reduction of the maximum lift coefficient caused by different equivalent sand-grain roughness heights and Reynolds number present similar behavior, they fall into different curve formats. </abstract>

Upper surface flaps commonly referred to as spoilers or drag brakes can increase maximum lift, and improve aerodynamic efficiency at high, near-stall angles of attack. This phenomenon was studied experimentally and computationally using a 0.307626 m chord length NACA 2412 airfoil in six different configurations, and one baseline clean configuration. A wind tunnel model was placed in the Ryerson Low Speed Wind Tunnel (atmospheric, closed-circuit, 3 ft × 3 ft test section) at a Reynold’s number of approximately 780,000 and a Mach number of 0.136. The wind tunnel study increased the lift coefficient by 0.393%-2.497% depending on the spoiler configuration. A spoiler of 10% chord length increased the maximum lift coefficient by 2.497 % when deflected 8º, by 2.110% when deflected 15º, and reduced the maximum lift coefficient by 2.783% when deflected 25º. A spoiler of 15% chord length produced smaller maximum lift coefficient gains; 0.393% when deflected 8º, by 1.760% when deflected 15º, and reduced the maximum lift coefficient by 4.475% when deflected 25º. Deflecting the spoiler increased the stall angle between 37.658% and 87.544% when compared with the clean configuration. The drag coefficient of spoiler configurations was lower than the clean configuration at angles of attack above 18º. The combination of the increased lift and reduced drag at angles of attack above 18º created by the spoiler configurations resulted in a higher aerodynamic efficiency than the clean configuration case. A 10% chord length spoiler deflected at 8º produced the highest aerodynamic efficiency gains. At low angles of attack, the computational study produced consistently higher lift coefficients compared with the wind tunnel experiment. The lift-slope was consistent with the wind tunnel experiment lift-slope. The spoiler airfoil stall behaviour was inconsistent with the results from the wind tunnel experiment. The drag coefficient results were consistent with the wind tunnel experiment at low angles of attack. However, the spoiler equipped airfoils did not reduce drag at high angles of attack. Therefore, the computational model was not valid for the spoiler configurations at high angles of attack.

Effects of Chordwise Flexibility on the Aerodynamic Performance of a 3D Flapping Wing

Modern wind energy industry is highly under the influence of the advances in aerospace and has borrowed and adopted many new and innovative concepts from that industry. With the aim of exploring potential benefits of flapping blades in wind turbines, this paper presents the preliminary results of the numerical simulation of a flapping-pitching rectangular flat plate in a uniform air flow. Various combinations of flapping amplitude, flapping frequency and pitching amplitude are analysed and their effect on the instantaneous and maximum lift coefficient is presented. The change in the flapping frequency and amplitude were shown to have considerable effect on the lift coefficient. It can be deduced from the results that the lift coefficient is influenced by the flapping frequency and flapping amplitude combination. For the size and geometry presented in this study, the lift coefficient is most affected by the flapping amplitude when compared to the flapping frequency. The results indicate that the pitching amplitude initially enhances the lift coefficient. However, excessive pitching amplitude results in low lift coefficient.

∗† ‡ Force measurements over a range of non-slender delta wings have demonstrated the ability of a flexible wing to enhance lift and delay stall compared with a rigid wing of similar geometry. The work has extended the results of a recent study to include delta wings with a range of sweep angles. It has been shown that the greatest lift enhancement was observed over the wing with smallest leading edge sweep. Additionally for the only slender wing considered no lift enhancement was observed indicating that it is a feature of the fundamentally different flow that occurs over low-sweep wings that is responsible for the phenomenon. The variation of RMS lift force and rolling moment coefficients with incidence for all the wings concerned suggest that the wings are vibrating in an antisymmetric structural mode in the lift enhancement region. This supports previous evidence for the 50° wing that suggests the same. Further, experiments of a half-wing model suggest that this wing does not undergo lift enhancement, and as such it may be that the antisymmetric vibration is a necessary condition for the lift enhancement to exist. Wing-tip deflections for some of the wings were also measured, which showed that the slender wing undergoes much higher time-averaged deflections than the non-slender wings, although it also experiences a much lower level of buffet. PIV and LDV measurements have demonstrated the striking difference between the surface flows over the flexible and rigid wings in the post-stall region at the same incidence. Implementing flexibility on a low sweep wing results in continued reattachment of the shear layer to much higher incidences than would otherwise be observed over a rigid wing.

A coupled CFD-CSD solver is used to simulate the aerodynamics of flexible flapping wings for MAV applications. The CFD solver is a compressible RANS solver developed at University of Maryland, called OVERTURNS. The open source general-purpose multibody dynamics solver MBDyn is used as the structural solver after an extension to take into account geometrically nonlinear shells, making it possible to analyze plates/membranes with large deformations. Validation of the two codes for flapping wings/plates is carried out independently. The solvers are then coupled and validated against prior experiments and predictions of a simplified configuration involving a spanwise flexible wing in pure plunge. In-house experiments are also carried out on MAV wings in an open jet wind tunnel. Rigid wings are tested for pure flap and flap plus pitch kinematics. Flexible wings with root flap actuation are also tested in the wind tunnel and experiments are carried out to structurally characterize them. CFD and CFD-CSD analyses are carried out on the rigid and flexible wings, respectively. Good correlation is observed between the experimental and analytical results. Further, the analysis is used to study the flow physics in greater detail. A high fidelity, well validated CFD-CSD solver is developed which may help in designing the next generation flapping wing MAVs.

An experimental study on flapping wing flexibility in hovering flight has been conducted to investigate the wing flexibility for insect-inspired flapping Micro Aerial Vehicles (MAVs). Hawkmoth-like wing models, derived from Manduca sexta, were made of Polycarbonate (PC) sheet with a spanwise length of 200 mm and an aspect ratio of 6.18. For the distributions of wing flexibility, the wing thickness was selected as the design variable: rigid wing (3 mm-thick) and flexible wings (2, 1, 0.8, 0.5, 0.35, 0.2, and 0.1 mm-thick). In the experiment, the wing models were constrained to the symmetrical and sinusoidal flapping motions with sweeping and rotating amplitudes of 120° and 90° in water tank with size of 3.5 m×1.0 m×1.1 m. Aerodynamic force and flow structures for flapping the wing were measured using a six-axis force/torque sensor and a high speed camera with a laser using Digital Particle Image Velocimetry (DPIV). To compare the flow structures of flexible wings with rigid wing, they were captured at the same chordwise cross-section as the rigid wing, 50% of wing length. Based on the experimental results, vortices and aerodynamic force. Consequently, the wing with thickness of 0.8 mm has better aerodynamic characteristics than other wings in hovering flight. This finding will be instrumental in identifying the range of wing flexibilities that improves the aerodynamic efficiency for the development of insect-inspired flapping MAVs.

A computational study has been conducted on two wings of aspect ratios 1.244 and 1.865, each having 65-deg leading edge sweep angles, to determine the effects of nonplanar winglets at supersonic Mach numbers. A design Mach number of 1.62 was selected. The winglets studied were parametrically varied in alignment, length, sweep, camber, and thickness to determine the effects of winglet geometry on predicted performance. For the computational analysis, an existing Euler code that employed a marching technique was used. The results indicated that the possibility existed for wing-winglet geometries to equal the performance of wing-alone bodies in supersonic flows with both bodies having the same semispan length. The performance parameters of main interest were the lift-to-pressure drag ratio and the pressure drag coefficient as functions of lift coefficient. The lift coefficient range for this study was from -0.20 to 0.70 with emphasis on the range of 0.10 to 0.22.