scholarly journals Effect of Flexible Flaps on Lift and Drag of Laminar Profile Flow

Energies ◽  
2020 ◽  
Vol 13 (5) ◽  
pp. 1077
Author(s):  
Artur Reiswich ◽  
Max Finster ◽  
Martin Heinrich ◽  
Rüdiger Schwarze
Keyword(s):  
Wind Tunnel ◽  
Critical Angle ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Force Balance ◽  
Airfoil Surface ◽  
Pressure Drag ◽  
Positive Effects ◽  
Trailing Edges ◽  
Lift And Drag

Experiments with elastic flaps applied on a common airfoil profile were performed to investigate positive effects on lift and drag coefficients. An NACA0020 profile was mounted on a force balance and placed in a wind tunnel. Elastic flaps were attached in rows at different positions on the upper profile surface. The Reynolds number of the flow based on the chord length of the profile is about 2 × 10 5 . The angle of attack is varied to identify the pre- and post-stall effects of the flaps. Polar diagrams are presented for different flap configurations to compare the effects of the flaps. The results showed that flaps generally increase the drag coefficient due to the additional skin friction and pressure drag. Furthermore, a significant increase of lift in the stall region was observed. The highest efficiency was obtained for the configuration with flaps at the leading and trailing edges of the profile. In this case, the critical angle was delayed and lift was increased in pre- and post-stall regions. This flap configuration was used in a gust simulation in the wind tunnel to model unsteady incoming flow at a critical angle of attack. This investigation showed that the flow separation at the critical angle was prevented. Additionally, smoke–wire experiments were performed for the stall region in order to visualize the flow around the airfoil. The averaged flow field results showed that the leading-edge flaps lean the flow more towards the airfoil surface and reduce the size of the separated region. This reduction improves the airfoil performance in the deep stall region.

Alula-Inspired Leading Edge Device for Low Reynolds Number Flight

2016 ◽  
Author(s):  
Boris A. Mandadzhiev ◽  
Michael K. Lynch ◽  
Leonardo P. Chamorro ◽  
Aimy A. Wissa
Keyword(s):  
Reynolds Number ◽  
Wind Tunnel ◽  
Drag Coefficient ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Tip Vortex ◽  
Relative Angle ◽  
Lift And Drag

Robust and predictable aerodynamic performance of unmanned aerial vehicles at the limits of their design envelope is critical for safety and mission adaptability. In order for a fixed wing aircraft to maintain the lift necessary for sustained flight at very low speeds and large angles of attack (AoA), the wing shape has to change. This is often achieved by using deployable aerodynamic surfaces, such as flaps or slats, from the wing leading or trailing edges. In nature, one such device is a feathered structure on birds’ wings called the alula. The span of the alula is 5% to 20% of the wing and is attached to the first digit of the wing. The goal of the current study is to understand the aerodynamic effects of the alula on wing performance. A series of wind tunnel experiments are performed to quantify the effect of various alula deployment parameters on the aerodynamic performance of a cambered airfoil (S1223). A full wind tunnel span wing, with a single alula located at the wing mid-span is tested under uniform low-turbulence flow at three Reynolds numbers, Re = 85,000, 106,00 and 146,000. An experimental matrix is developed to find the range of effectiveness of an alula-type device. The alula relative angle of attack measured measured from the mean chord of the airfoil is varied to modulate tip-vortex strength, while the alula deflection is varied to modulate the distance of the tip vortex to the wing surface. Lift and drag forces were measured using a six axis force transducer. The lift and drag coefficients showed the greatest sensitivity to the the alula relative angle of attack, increasing the normalized lift coefficient by as much as 80%. Improvements in lift are strongly correlated to higher alula angle, with β = 0° – 5°, while reduction in the drag coefficient is observed with higher alula tip deflection ratios and lower β angles. Results show that, as the wing angle of attack and Reynolds number are increased, the overall lift co-efficient improvement is diminished while the reduction in drag coefficient is higher.

scholarly journals Design and Validation of a New Morphing Camber System by Testing in the Price—Païdoussis Subsonic Wind Tunnel

Aerospace ◽  
2020 ◽  
Vol 7 (3) ◽  
pp. 23 ◽  
Author(s):  
David Communier ◽  
Ruxandra Mihaela Botez ◽  
Tony Wong
Keyword(s):  
Wind Tunnel ◽  
Unmanned Aerial Vehicle ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Trailing Edge ◽  
Subsonic Wind Tunnel ◽  
Aerial Vehicle ◽  
Stall Angle

This paper presents the design and wind tunnel testing of a morphing camber system and an estimation of performances on an unmanned aerial vehicle. The morphing camber system is a combination of two subsystems: the morphing trailing edge and the morphing leading edge. Results of the present study show that the aerodynamics effects of the two subsystems are combined, without interfering with each other on the wing. The morphing camber system acts only on the lift coefficient at a 0° angle of attack when morphing the trailing edge, and only on the stall angle when morphing the leading edge. The behavior of the aerodynamics performances from the MTE and the MLE should allow individual control of the morphing camber trailing and leading edges. The estimation of the performances of the morphing camber on an unmanned aerial vehicle indicates that the morphing of the camber allows a drag reduction. This result is due to the smaller angle of attack needed for an unmanned aerial vehicle equipped with the morphing camber system than an unmanned aerial vehicle equipped with classical aileron. In the case study, the morphing camber system was found to allow a reduction of the drag when the lift coefficient was higher than 0.48.

Linearized Theory for Supercavitating Hydrofoils With Spoiler-Flaps

1962 ◽  
Vol 6 (04) ◽  
pp. 1-9 ◽  
Author(s):  
C. F. Chen
Keyword(s):  
Leading Edge ◽  
Hydrodynamic Characteristics ◽  
Arbitrary Length ◽  
Infinite Depth ◽  
Trailing Edges ◽  
Linearized Problem ◽  
Linearized Theory ◽  
Hinge Position ◽  
Trailing Edge Flap ◽  
Lift And Drag

The linearized problem of flow past a flat plate with a deflected upper flap is solved including the free-surface effects. The flow is assumed to detach at the trailing edges of the flap and of the plate, leaving an infinite cavity of constant pressure behind. The results are applicable to flaps of arbitrary length, hinged at an arbitrary distance from the leading edge and at arbitrary depth. Numerical results are presented to show the effect of depth on the lift and drag coefficients of a trailing-edge flap whose chord is 30 percent of the chord of the plate, the effect of the hinge position of such a flap on the hydrodynamic characteristics of a plate at infinite depth, and the pressure distribution on a flapped plate at infinite depth. The principal results obtained are as follows: As the depth decreases, the flap becomes less effective and less efficient; as the hinge position is moved rearward, the flap becomes more effective and more efficient.

scholarly journals Genetic Algorithm Based Optimization of Wing Rotation in Hover

Fluids ◽  
2018 ◽  
Vol 3 (3) ◽  
pp. 59 ◽  
Author(s):  
Alexander Gehrke ◽  
Guillaume Guyon-Crozier ◽  
Karen Mulleners
Keyword(s):  
Genetic Algorithm ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Airfoil Surface ◽  
Pitching Motion ◽  
Leading Edge Vortex ◽  
Final Population ◽  
Edge Vortex ◽  
Effective Angle

The pitching kinematics of an experimental hovering flapping wing setup are optimized by means of a genetic algorithm. The pitching kinematics of the setup are parameterized with seven degrees of freedom to allow for complex non-linear and non-harmonic pitching motions. Two optimization objectives are considered. The first objective is maximum stroke average efficiency, and the second objective is maximum stroke average lift. The solutions for both optimization scenarios converge within less than 30 generations based on the evaluation of their fitness. The pitching kinematics of the best individual of the initial and final population closely resemble each other for both optimization scenarios, but the optimal kinematics differ substantially between the two scenarios. The most efficient pitching motion is smoother and closer to a sinusoidal pitching motion, whereas the highest lift-generating pitching motion has sharper edges and is closer to a trapezoidal motion. In both solutions, the rotation or pitching motion is advanced with respect to the sinusoidal stroke motion. Velocity field measurements at selected phases during the flapping motions highlight why the obtained solutions are optimal for the two different optimization objectives. The most efficient pitching motion is characterized by a nearly constant and relatively low effective angle of attack at the start of the half stroke, which supports the formation of a leading edge vortex close to the airfoil surface, which remains bound for most of the half stroke. The highest lift-generating pitching motion has a larger effective angle of attack, which leads to the generation of a stronger leading edge vortex and higher lift coefficient than in the efficiency optimized scenario.

Impact of Gurney Flaplike Strips on the Aerodynamic and Vortex Flow Characteristic of a Reverse Delta Wing

2016 ◽  
Vol 138 (6) ◽  
Author(s):  
T. Lee
Keyword(s):  
Vortex Flow ◽  
Delta Wing ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Force Balance ◽  
Vortex Control ◽  
Lift Curve ◽  
Lift And Drag ◽  
The Impact

The impact of Gurney flaplike strips, of different geometric configurations and heights, on the aerodynamic characteristics and the tip vortices generated by a reverse delta wing (RDW) was investigated via force-balance measurement and particle image velocimetry (PIV). The addition of side-edge strips (SESs) caused a leftward shift of the lift curve, resembling a conventional trailing-edge flap. The large lift increment overwhelmed the corresponding drag increase, thereby leading to an improved lift-to-drag ratio compared to the baseline wing. The lift and drag coefficients were also found to increase with the strip height. The SES-equipped wing also produced a strengthened vortex compared to its baseline wing counterpart. The leading-edge strips (LESs) were, however, found to persistently produce a greatly diffused vortex flow as well as a small-than-baseline-wing lift in the prestall α regime. The downward LES delivered a delayed stall and an increased maximum lift coefficient compared to the baseline wing. The LESs provide a potential wingtip vortex control alternative, while the SESs can enhance the aerodynamic performance of the RDW.

scholarly journals Aerodynamic Study of the NASA’s X-43A Hypersonic Aircraft

2020 ◽  
Vol 10 (22) ◽  
pp. 8211
Author(s):  
Àlex Navó ◽  
Josep M. Bergada
Keyword(s):  
Angle Of Attack ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Oblique Shock ◽  
Force Balance ◽  
Oblique Shock Wave ◽  
Vertical Force ◽  
Pitching Moment ◽  
Cfd Model ◽  
Hypersonic Aircraft

A 2D aerodynamic study of the NASA’s X-43A hypersonic aircraft is developed using two different approaches. The first one is analytical and based on the resolution of the oblique shock wave and Prandtl–Meyer expansion wave theories supported by an in-house program and considering a simplified aircraft’s design. The second approach involves the use of a Computational Fluid Dynamics (CFD) package, OpenFOAM and the real shape of the aircraft. The aerodynamic characteristics defined as the lift and drag coefficients, the aerodynamic efficiency and the pitching moment coefficient are calculated for different angles of attack. Evaluations are made for an incident Mach number of 7 and an altitude of 30 km. For both methodologies, the required angles of attack to achieve a Vertical Force Balance (VFB) and a completely zero pitching moment conditions are considered. In addition, an analysis to optimise the nose configuration of the aircraft is performed. The mass flow rate throughout the scramjet as a function of the angle of attack is also presented in the CFD model in addition to the pressure, density, temperature and Mach fields. Before presenting the corresponding results, a comparison between the aerodynamic coefficients in terms of the angle of attack of both models is carried out in order to properly validate the CFD model. The paper clarifies the requirements needed to make sure that both oblique shock waves originating from the leading edge meet just at the scramjet inlet clarifying the advantages of fulfilling such condition.

Vortex Generators Contribution to the Enhancement of the Aerodynamic Performances

2014 ◽  
Vol 950 ◽  
pp. 268-274
Author(s):  
Hocine Tebbiche ◽  
Mohamed S. Boutoudj
Keyword(s):  
Wind Tunnel ◽  
Flow Control ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Vortex Generators ◽  
Drag Coefficients ◽  
Passive Devices ◽  
Aerodynamic Performances ◽  
Lift And Drag ◽  
Naca 0015

This study interest flow control using a new vortex generators (VGs) shape with counter-rotating vortices, obtained by adding a new element to a configuration mostly investigated. The experiments were performed in the aim to determine the VGs answer when placed on the suction face at 10% from the leading edge of an airfoil Naca 0015 in order to improve the lift and drag coefficients. The investigations were accomplished in wind tunnel for two Reynolds numbers and geometrical vortex generators configurations. The obtained results are analyzed according to several parameters such as the VG height, the space between the same VG pair and the additional factor. The results show a profit brought by the passive devices estimated at about 28% of the CL/Cd ratio.

scholarly journals Studies Based on CFD Behaviour of Aerofoil and Regression Analysis

2021 ◽  
pp. 55-68
Author(s):  
Satish Geeri ◽  
Sambhu Prasad Surapaneni ◽  
Jithendra Sai Raja Chada ◽  
Akhil Yuvaraj Manda
Keyword(s):  
Regression Analysis ◽  
Surface Area ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Surface Modifications ◽  
Active Surface ◽  
Attack Angle ◽  
Inlet Velocity ◽  
Optimum Model ◽  
Lift And Drag

The performance of an aerofoil depends upon the angle of attack, leading-edge radius, surface modifications, etc. The aerofoil which has a broader range of attack angle and surface area is responsible for the upliftment in the performance of the aerofoil. The present work deals with the evaluation of the aerofoil spread with dimples over the active surface. The positions and area of spread are modified accordingly and evaluated for the velocity and pressure lineation. The aerofoil with 30% dimples over the active surface is found to possess higher values for the required intents of velocity and pressure at an inlet velocity of 9 m/s. The optimum model with better lineation values is further evaluated for the co-efficient of lift and drag to propose the best design. The best result is obtained at an aerofoil of NACA 8412 series with 30% dimples extension at the rear end placed at 15° angle of attack and the regression analysis is done for the coefficient of lift values.

Transonic flow of dense gases around an airfoil with a parabolic nose

1997 ◽  
Vol 346 ◽  
pp. 1-21 ◽  
Author(s):  
ZVI RUSAK ◽  
CHUN-WEI WANG
Keyword(s):  
Stagnation Point ◽  
Angle Of Attack ◽  
Outer Region ◽  
Leading Edge ◽  
Critical Conditions ◽  
Nonlinearity Parameter ◽  
Airfoil Surface ◽  
Dense Gases ◽  
Outer Expansion ◽  
Inner Region

Transonic potential flow of dense gases of retrograde type around the leading edge of a thin airfoil with a parabolic nose is studied. The analysis follows the approach of Rusak (1993) for a perfect gas. Asymptotic expansions of the velocity potential function are constructed in terms of the airfoil thickness ratio in an outer region around the airfoil and in an inner region near the nose. The outer expansion consists of the transonic small-disturbance theory for dense gases, where a leading-edge singularity appears. Analytical expressions are given for this singularity by constructing similarity solutions of the governing nonlinear equation. The inner expansion accounts for the flow around the nose, where a stagnation point exists. A boundary value problem is formulated in the inner region for the solution of an oncoming uniform sonic flow with zero values of the fundamental derivative of gasdynamics (Γ=0) and the second nonlinearity parameter (Λ=0) around a parabola at zero angle of attack. The numerical solution of the inner problem results in a symmetric flow around the nose. The outer and inner expansions are matched asymptotically resulting in a uniformly valid solution on the entire airfoil surface. In the leading terms, the flow around the nose is symmetric and the stagnation point is located at the leading edge for every transonic Mach number, and small values of Γ and Λ of the oncoming flow and any shape and small angle of attack of the airfoil. Furthermore, analysis of the inner region in the immediate neighbourhood of the stagnation point reveals that the flow is purely subsonic, approaching critical conditions in the limit of large (scaled) distances, which excludes the formation of shock discontinuities in the nose region.

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