High-Lift Mechanism of a Bionic Slat

2013 ◽  
Vol 461 ◽  
pp. 220-229
Author(s):  
Chang Jiang Ge ◽  
Mei Chen Ge
Keyword(s):  
Flow Field ◽  
Pressure Gradient ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Adverse Pressure Gradient ◽  
Element Model ◽  
Broadband Noise ◽  
High Lift ◽  
Edge Separation

To avoid broadband noise from a slat cove, the deployed slat contour is usually modified by filling cove, but the design is sensitive to aerodynamic performance. In the paper, a bionic slat without a cove is built on the basis of a bionic airfoil (i.e. stowed bionic multi-element airfoil), which is extracted from a long-eared owl wing. The quasi-two-dimensional models with a deployed bionic slat and a stowed bionic slat are manufactured by rapid manufacturing and prototyping system, respectively, and measured in a low-turbulence wind tunnel. The results are used to characterize high-lift effect: the lift coefficients of the model with a stowed slat are larger at less than 4°angle of attack, but the model with a deployed slat has the larger lift coefficients at greater than 4°angle of attack. Furthermore, the deployed bionic slat can increase stall angle and maximum lift coefficient, but also delay the decline of the lift coefficient curve slope meaning that the leading-edge separation is postponed within a certain range of angle of attack. At the same time, the flow field around the models is visualized by smoke wire method. The leading-edge separation of the model with a stowed slat is shown at low Reynolds number and angle of attack. However, the finding does not occur in the flow field of the model with a deployed slat at the same conditions, probably because the gap between the bionic slat and the main wing results in favorable pressure gradient, the deployed bionic slat decreases the peak of adverse pressure gradient by increasing the chord of the bionic multi-element model, and the bionic slat wake excites transition to the boundary layer on upper surface of the main wing. This superiority may be used as reference in the design of the leading-edge slat without a cove.

scholarly journals Study of Vortex Generator Effect on Airfoil Aerodynamics Using the Computational Fluids Dynamics Method

2020 ◽  
Vol 3 (2) ◽  
pp. 23
Author(s):  
Siti Aisyah Ayudia ◽  
Artoto Arkundato ◽  
Lutfi Rohman
Keyword(s):  
Boundary Layer ◽  
Lift Force ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Vortex Generator ◽  
Leading Edge ◽  
Adverse Pressure Gradient ◽  
Separation Flow ◽  
Main Thing ◽  
Dynamics Method

The lift force is one of the important factors in supporting the aircraft flying capabilities. The airplane has a section called the aircraft wing. In particular, the wing section of aircraft is called the airfoil. One of the efforts to increase the lift force is to make the flow of air fluid at the top of the airfoil more turbulent. Turbulent flow can attract momentum from the boundary layer, the result of this momentum transfer has energy that is more resistant to the adverse pressure gradient which can trigger the flow separation. Efforts that can be made to reduce separation flow and increase lift force are the addition of a turbulent generator on the upper surface of the airfoil, one type of turbulent generator is a vortex generator, a vortex generator can accelerate the transition from the laminar boundary layer to the turbulent boundary layer. This study was conducted with the aim of knowing the effect of the vortex generator on the aerodynamics of NACA-4412 using the computational fluid dynamics method. The main thing that will be investigated is the effect of the straight type vortex generator application on the lift coefficient, by comparing the plain airfoil and airfoil that has been applied to the vortex generator to vary the angle of attack. The variation of the angles of attack are 0º, 5º, 10º, 15º and the placement of the vortex generator is 24% of the leading edge. The results obtained that the lift coefficient changes with increasing angle of attack and the application of a vortex generator to an airfoil can increase the lift coefficient than a plain airfoil. The optimum increase in lift coefficient is at the angle of attack of 5º as much as 13%.

scholarly journals A STUDY OF HIGH LIFT AERODYNAMIC DEVICES ON COMMERCIAL AIRCRAFTS

Aviation ◽  
2020 ◽  
Vol 24 (3) ◽  
pp. 123-136
Author(s):  
Swamy Naidu Venkata Neigapula ◽  
Satya Prasad Maddula ◽  
Vasishta Bhargava Nukala
Keyword(s):  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Trailing Edge ◽  
High Lift ◽  
Aircraft Wings ◽  
Surface Pressure Distribution ◽  
Trailing Edge Flaps ◽  
Stall Angle ◽  
Kinematic Mechanisms

Aerodynamic performance of aircraft wings vary with flight path conditions and depend on efficiency of high lift systems. In this work, a study on high lift devices and mechanisms that aim to increase maximum lift coefficient and reduce drag on commercial aircraft wings is discussed. Typically, such extensions are provided to main airfoil along span wise direction of wing and can increase lift coefficient by more than 100% during operation. Increasing the no of trailing edge flaps in chord wise direction could result in 100% increment in lift coefficient at a given angle of attack but leading edge slats improve lift by delaying the flow separation near stall angle of attack. Different combinations of trailing edge flaps used by Airbus, Boeing and McDonnel Douglas manufacturers are explained along with kinematic mechanisms to deploy them. The surface pressure distribution for 30P30N airfoil is evaluated using 2D vortex panel method and effects of chord wise boundary layer flow transitions on aerodynamic lift generation is discussed. The results showed better agreements with experiment data for high Reynolds number (9 million) flow conditions near stall angle of attack.

scholarly journals Effect of airfoil distance to water surface on static stall

2020 ◽  
Vol 14 (1) ◽  
pp. 6526-6537
Author(s):  
A. Yeganeh ◽  
Mohammad Hassan Djavareshkian ◽  
E. Esmaeil
Keyword(s):  
Pressure Gradient ◽  
Water Surface ◽  
Stokes Equations ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Adverse Pressure Gradient ◽  
Free Surface Flow ◽  
Surface Flow ◽  
Navier Stokes ◽  
Navier Stokes Equations

In this study, viscous, turbulent, and steady flow around an airfoil near the water surface has been simulated through a numerical method. In this simulation, Navier-Stokes equations have been solved using the finite volume method with a discretized second-order accuracy and PIMPLE algorithm. The Volume of Fraction (VOF) method has been employed to predict the free surface flow. A part of the simulation results has been validated through numerical and experimental data. Besides considering the style of flow separation in the angles of numerous attacks and airfoil static stall near the surface of the water. For this purpose, the airfoil simulation has been processed airfoil in the 68,000 Reynolds number, angle of attack of 2.5 to 11 degree and different distances from the water surface ( h/c = 0.5, 1,  ). In a larger angle of attacks, flow is initially separated from the leading edge of the surface, and then it attaches to the surface at a lower point. This reattachment leads to an increase in adverse pressure gradient and the formation of a larger separation in the downstream of the airfoil. The pressure gradient dramatically increases, and the flow gets separated from the upstream of the airfoil. Upon lowering distance from the surface, static stall takes place at a higher point and a lower angle of attack, respectively.

Measurements in adverse-pressure-gradient turbulent boundary layers with a step change in surface roughness

1975 ◽  
Vol 70 (3) ◽  
pp. 573-593 ◽  
Author(s):  
W. H. Schofield
Keyword(s):  
Surface Roughness ◽  
Pressure Gradient ◽  
Boundary Layers ◽  
Leading Edge ◽  
Adverse Pressure Gradient ◽  
Turbulent Boundary Layers ◽  
Step Change ◽  
Internal Layer ◽  
Mean Velocity ◽  
Law Of The Wall

The response of turbulent boundary layers to sudden changes in surface roughness under adverse-pressure-gradient conditions has been studied experimentally. The roughness used was in the ‘d’ type array of Perry, Schofield & Joubert (1969). Two cases of a rough-to-smooth change in surface roughness were considered in the same arbitrary adverse pressure gradient. The two cases differed in the distance of the surface discontinuity from the leading edge and gave two sets of flow conditions for the establishment and growth of the internal layer which develops downstream from a change in surface roughness. These conditions were in turn different from those in the zero-pressure-gradient experiments of Antonia & Luxton. The results suggest that the growth of the new internal layer depends solely on the new conditions at the wall and scales with the local roughness length of that wall. Mean velocity profiles in the region after the step change in roughness were accurately described by Coles’ law of the wall-law of the wake combination, which contrasts with the zero-pressure-gradient results of Antonia & Luxton. The skin-friction coefficient after the step change in roughness did not overshoot the equilibrium distribution but made a slow adjustment downstream of the step. Comparisons of mean profiles indicate that similar defect profile shapes are produced in layers with arbitrary adverse pressure gradients at positions where the values of Clauser's equilibrium parameter β (= δ*τ−10dp/dx) are similar, provided that the pressure-gradient history and local values of the pressure gradient are also similar.

Computational Design and Experimental Evaluation of Using a Leading Edge Fillet on a Gas Turbine Vane

2001 ◽  
Author(s):  
G. A. Zess ◽  
K. A. Thole
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Total Pressure ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Computational Design ◽  
Horseshoe Vortex

With the desire for increased power output for a gas turbine engine comes the continual push to achieve higher turbine inlet temperatures. Higher temperatures result in large thermal and mechanical stresses particularly along the nozzle guide vane. One critical region along a vane is the leading edge-endwall juncture. Based on the assumption that the approaching flow to this juncture is similar to a two-dimensional boundary layer, previous studies have shown that a horseshoe vortex forms. This vortex forms because of a radial total pressure gradient from the approaching boundary layer. This paper documents the computational design and experimental validation of a fillet placed at the leading edge-endwall juncture of a guide vane to eliminate the horseshoe vortex. The fillet design effectively accelerated the incoming boundary layer thereby mitigating the effect of the total pressure gradient. To verify the CFD studies used to design the leading edge fillet, flow field measurements were performed in a large-scale, linear, vane cascade. The flow field measurements were performed with a laser Doppler velocimeter in four planes orientated orthogonal to the vane. Good agreement between the CFD predictions and the experimental measurements verified the effectiveness of the leading edge fillet at eliminating the horseshoe vortex. The flowfield results showed that the turbulent kinetic energy levels were significantly reduced in the endwall region because of the absence of the unsteady horseshoe vortex.

scholarly journals On the lift-optimal aspect ratio of a revolving wing at low Reynolds number

2018 ◽  
Vol 15 (143) ◽  
pp. 20170933 ◽  
Author(s):  
T. Jardin ◽  
T. Colonius
Keyword(s):  
Aspect Ratio ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Rossby Number ◽  
Subsequent Work ◽  
High Lift ◽  
Low Aspect Ratio ◽  
Axis Of Rotation ◽  
Edge Vortex ◽  
Convergent Solution

Lentink & Dickinson (2009 J. Exp. Biol. 212 , 2705–2719. ( doi:10.1242/jeb.022269 )) showed that rotational acceleration stabilized the leading-edge vortex on revolving, low aspect ratio (AR) wings and hypothesized that a Rossby number of around 3, which is achieved during each half-stroke for a variety of hovering insects, seeds and birds, represents a convergent high-lift solution across a range of scales in nature. Subsequent work has verified that, in particular, the Coriolis acceleration plays a key role in LEV stabilization. Implicit in these results is that there exists an optimal AR for wings revolving about their root, because it is otherwise unclear why, apart from possible morphological reasons, the convergent solution would not occur for an even lower Rossby number. We perform direct numerical simulations of the flow past revolving wings where we vary the AR and Rossby numbers independently by displacing the wing root from the axis of rotation. We show that the optimal lift coefficient represents a compromise between competing trends with competing time scales where the coefficient of lift increases monotonically with AR, holding Rossby number constant, but decreases monotonically with Rossby number, when holding AR constant. For wings revolving about their root, this favours wings of AR between 3 and 4.

scholarly journals An Inverse Design Method for Airfoils Based on Pressure Gradient Distribution

Energies ◽  
2020 ◽  
Vol 13 (13) ◽  
pp. 3400
Author(s):  
Yufei Zhang ◽  
Chongyang Yan ◽  
Haixin Chen
Keyword(s):  
Pressure Gradient ◽  
Design Method ◽  
Lift Coefficient ◽  
Pressure Coefficient ◽  
Poor Performance ◽  
Adverse Pressure Gradient ◽  
Inverse Design ◽  
Pressure Gradients ◽  
Gradient Distribution ◽  
Inverse Design Method

An airfoil inverse design method is proposed by using the pressure gradient distribution as the design target. The adjoint method is used to compute the derivatives of the design target. A combination of the weighted drag coefficient and the target dimensionless pressure gradient is applied as the optimization objective, while the lift coefficient is considered as a constraint. The advantage of this method is that the designer can sketch a rough expectation of the pressure distribution pattern rather than a precise pressure coefficient under a certain lift coefficient and Mach number, which can greatly reduce the design iteration in the initial stage of the design process. Multiple solutions can be obtained under different objective weights. The feasibility of the method is validated by a supercritical airfoil and a supercritical natural laminar flow airfoil, which are designed based on the target pressure gradients on the airfoils. Eight supercritical airfoils are designed under different upper surface pressure gradients. The drag creep and drag divergence characteristics of the airfoils are numerically tested. The shockfree airfoil demonstrates poor performance because of a high suction peak and the double-shock phenomenon. The adverse pressure gradient on the upper surface before the shockwave needs to be less than 0.2 to maintain both good drag creep and drag divergence characteristics.

Aeroelastic Analysis of Membrane Micro Air Vehicles

2009 ◽  
Author(s):  
Peter J. Attar ◽  
Raymond E. Gordnier ◽  
Jordan W. Johnston ◽  
William A. Romberg ◽  
Ramkumar N. Parthasarathy
Keyword(s):  
Strouhal Number ◽  
Limit Cycle ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Micro Air Vehicles ◽  
Element Model ◽  
Wing Membrane ◽  
Reduced Frequency ◽  
Membrane Wing ◽  
Air Vehicles

The fluid and structural response of two different membrane wing Micro Air Vehicles is studied through computation and experiment. A (three) batten-reinforced fixed wing membrane micro air vehicle is used to determine the effect of membrane prestrain and fixed angle of attack on flutter and limit cycle behavior of fixed wing membrane Micro Air Vehicles. For each configuration tested, flutter and subsequent limit cycle oscillations are measured in wind tunnel tests and predicted using an aeroelastic computational model consisting of a nonlinear finite element model coupled to a vortex lattice solution of the Laplace equation and boundary conditions. Correlation between the predicted and measured onset of limit cycle oscillation is good as is the prediction of the amplitude of the limit cycle at the trailing edge of the lower membrane. A direct correlation between levels of strain and the phase of the membranes during the limit cycle is found in the computation and thought to also occur in the experiment. The second membrane wing micro air vehicle configuration is that of a plunging membrane airfoil model. This model is studied computationally using a sixth-order finite difference solution of the Navier-Stokes equations coupled to a nonlinear string finite element model. The effect, on the structural and fluid response, of plunging Strouhal number, reduced frequency and static angle of attack is examined. At two degree angle of attack, and Strouhal number of 0.2, the effect of increasing the plunging reduced frequency is to decrease the sectional lift coefficient and increase the sectional drag coefficient. At this angle of attack, minimal change in the sectional lift coefficient is found when increasing from a Strouhal number of 0.2 to 0.5 at reduced frequencies of 0.5 and 5.903, the lowest and highest values of this parameter which are studied in this work. For this angle of attack the maximum change which occurs when increasing the Strouhal number from 0.2 to 0.5 is at a reduced frequency of 1.5. When the effect of angle of attack is studied, it is found that at a Strouhal number of 0.5 and reduced frequency of 1.5 the plunging flexible model demonstrates improved lift characteristics over the fixed flexible airfoil case. The greatest improvement occurs at an angle of attack of 2 degrees followed by 10 degrees and then 6 degrees. Finally the effect on the flow characteristics of airfoil flexibility is investigated by increasing the membrane pre-strain from a nominal value of 5 percent to that of 20 percent. This increase in pre-strain results in a reduced value of sectional lift coefficient as compared the 5 percent pre-strain case at the same fixed angle of attack, Strouhal number and reduced frequency.

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.

scholarly journals Numerical investigation of the low-frequency flow oscillation over a NACA-0012 aerofoil at the inception of stall

2019 ◽  
Vol 11 ◽  
pp. 175682931983368 ◽  
Author(s):  
Yasir A ElAwad ◽  
Eltayeb M ElJack
Keyword(s):  
Experimental Data ◽  
Flow Field ◽  
Separation Bubble ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Low Frequency ◽  
Laminar Separation Bubble ◽  
Flow Oscillation ◽  
Laminar Separation ◽  
Naca 0012

High-fidelity large eddy simulation is carried out for the flow field around a NACA-0012 aerofoil at Reynolds number of [Formula: see text], Mach number of 0.4, and various angles of attack around the onset of stall. The laminar separation bubble is formed on the suction surface of the aerofoil and is constituted by the reattached shear layer. At these conditions, the laminar separation bubble is unstable and switches between a short bubble and an open bubble. The instability of the laminar separation bubble triggers a low-frequency flow oscillation. The aerodynamic coefficients oscillate accordingly at a low frequency. The lift and the drag coefficients compare very well to recent high-accuracy experimental data, and the lift leads the drag by a phase shift of [Formula: see text]. The mean lift coefficient peaks at the angle of attack of [Formula: see text], in total agreement with the experimental data. The spectra of the lift coefficient does not show a significant low-frequency peak at angles of attack lower than or equal the stall angle of attack ([Formula: see text]). At higher angles of attack, the spectra show two low-frequency peaks and the low-frequency flow oscillation is fully developed at the angle of attack of [Formula: see text]. The behaviour of the flow-field and changes in the turbulent kinetic energy over one low-frequency flow oscillation cycle are described qualitatively.

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