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 Karakteristik Wake Area Akibat Efek Penggunaan Vortex Generator di Belakang Wing Airfoil Naca 43018

2019 ◽  
Vol 4 (1) ◽  
pp. 55-63
Author(s):  
Setyo Hariyadi S.P ◽  
Wawan Aries Widodo
Keyword(s):  
Boundary Layer ◽  
Turbulent Boundary Layer ◽  
Flat Plate ◽  
Drag Force ◽  
Lift Force ◽  
Vortex Generator ◽  
Leading Edge ◽  
O 15 ◽  
O 19 ◽  
Wing Airfoil

Pada aliran yang melintasi suatu airfoil terdapat fenomena separasi, yakni ketika momentum aliran sudah tidak mampu lagi mengatasi adverse pressure gradien. Selanjutnya separasi ini akan diikuti dengan timbulnya daerah wake pada daerah di belakang airfoil yang mengakibatkan naiknya drag force dan menurunnya lift force. Untuk mengurangi hal tersebut maka vortex generator diletakkan pada sisi atas airfoil untuk mempercepat terbentuknya turbulent boundary layer sehingga dapat menunda separasi dan memperkecil daerah wake. Efektivitas dari vortex generator dipengaruhi oleh penempatan, ketinggian, dan interval antar vortex generator. Untuk mendapatkan hasil yang optimal, drag yang dihasilkan oleh vortex generator itu sendiri harus dikurangi. Untuk itu profil dari vortex generator yang digunakan harus sedemikian rupa sehingga drag yang dihasilkan dapat dikurangi tanpa menurunkan performasi dari airfoil tersebut. Oleh karena itu, penelitian ini dilakukan untuk melihat pengaruh penambahan vortex generator terhadap unjuk kerja airfoil melalui metode eksperimen. Tujuan penelitian ini adalah membandingkan karakteristik aliran fluida plain wing dan dengan penambahan vortex generator. Profil vortex generator yang digunakan adalah flat plate vortex generator dengan konfigurasi straight dan ditempatkan pada x/c = 10% dan 20% arah chord line dari leading edge. Variasi yang digunakan adalah bilangan Reynolds (Re), sudut serang (α) dan peletakan vortex generator pada airfoil. Kecepatan freestream yang digunakan yaitu kecepatan 12 m/s atau Re = 7,65 x 105 dan kecepatan 17 m/s atau Re = 9 x 105, dan pada sudut serang (α) 0o, 3 o, 6 o, 9 o, 12 o, 15 o, 19 o, dan 20 o. Hasil penelitian ini menunjukkan bahwa terjadi peningkatan performansi dari airfoil NACA 43018 dengan penambahan vortex generator dibandingkan dengan tanpa vortex generator. Adanya vortex generator, mempercepat perubahan dari aliran laminar ke turbulen. Separasi dapat tertunda dengan adanya vortex generator.

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 Leading Edge Blowing to Mimic and Enhance the Serration Effects for Aerofoil

2021 ◽  
Vol 11 (6) ◽  
pp. 2593
Author(s):  
Yasir Al-Okbi ◽  
Tze Pei Chong ◽  
Oksana Stalnov
Keyword(s):  
Boundary Layer ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Boundary Layer Separation ◽  
Shear Instability ◽  
Vortical Flow ◽  
High Angle ◽  
High Angle Of Attack ◽  
Layer Separation ◽  
Aerodynamic Performances

Leading edge serration is now a well-established and effective passive control device for the reduction of turbulence–leading edge interaction noise, and for the suppression of boundary layer separation at high angle of attack. It is envisaged that leading edge blowing could produce the same mechanisms as those produced by a serrated leading edge to enhance the aeroacoustics and aerodynamic performances of aerofoil. Aeroacoustically, injection of mass airflow from the leading edge (against the incoming turbulent flow) can be an effective mechanism to decrease the turbulence intensity, and/or alter the stagnation point. According to classical theory on the aerofoil leading edge noise, there is a potential for the leading edge blowing to reduce the level of turbulence–leading edge interaction noise radiation. Aerodynamically, after the mixing between the injected air and the incoming flow, a shear instability is likely to be triggered owing to the different flow directions. The resulting vortical flow will then propagate along the main flow direction across the aerofoil surface. These vortical flows generated indirectly owing to the leading edge blowing could also be effective to mitigate boundary layer separation at high angle of attack. The objectives of this paper are to validate these hypotheses, and combine the serration and blowing together on the leading edge to harvest further improvement on the aeroacoustics and aerodynamic performances. Results presented in this paper strongly indicate that leading edge blowing, which is an active flow control method, can indeed mimic and even enhance the bio-inspired leading edge serration effectively.

Control of Separated Boundary-Layer on Subsonic Airfoil Using Vortex Generator Jet Devices

2006 ◽  
Author(s):  
A. Kourta ◽  
G. Petit ◽  
J. C. Courty ◽  
J. P. Rosenblum
Keyword(s):  
Boundary Layer ◽  
Friction Coefficient ◽  
Skin Friction ◽  
Three Dimensional ◽  
Vortex Generator ◽  
Leading Edge ◽  
Skin Friction Coefficient ◽  
Synthetic Jets ◽  
High Lift ◽  
Longitudinal Vortices

The control of subsonic high lift induced separation on airfoil may improve the flight envelope of current aircraft or even simplify the complex and heavy high-lift devices on commercial airframes. Until now, synthetic jets have proved a really interesting efficiency to delay or remove even leading-edge located separated areas on high-lift configuration but are not efficient for real scale aircrafts. In case of pressure-like separation (i.e. from trailing-edge), synthetic jets can be replaced by so the called “Vortex Generator Jets” which create strong longitudinal vortices that increase mixing in inner boundary layer and consequently the skin friction coefficient is increased to prevent separation. In this study, numerical simulations were undertaken on a generic three dimensional flat plate in order to quantify the effect of the longitudinal vortices on the natural skin friction coefficient. Both counter and co-rotative devices were tested at different exhaust velocities and distances between each others. Finally co-rotative vortex generators jets were tested on a three dimensional generic airfoil ONERA D. Results show a delay of the separation occurence but this solution does not seem to be as robust as synthetic jets. The study of jets spacing with respect to the efficiency of the devices shows a maximum for a given ratio of spacing to exhaust velocity.

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 Analysis of Wind Turbine Airfoil Aerodynamic Performance with Leading Edge Bump

2015 ◽  
Vol 2015 ◽  
pp. 1-8 ◽  
Author(s):  
Majid Asli ◽  
Behnam Mashhadi Gholamali ◽  
Abolghasem Mesgarpour Tousi
Keyword(s):  
Wind Turbine ◽  
Control Method ◽  
Flow Behavior ◽  
Renewable Energy Sources ◽  
Lift Coefficient ◽  
Turbine Blades ◽  
Vortex Generator ◽  
Leading Edge ◽  
Suction Side ◽  
Aerodynamic Performance

Aerodynamic performance improvement of wind turbine blade is the key process to improve wind turbine performance in electricity generated and energy conversion in renewable energy sources concept. The flow behavior on wind turbine blades profile and the relevant phenomena like stall can be improved by some modifications. In the present paper, Humpback Whales flippers leading edge protuberances model as a novel passive stall control method was investigated on S809 as a thick airfoil. The airfoil was numerically analyzed by CFD method in Reynolds number of 106and aerodynamic coefficients in static angle of attacks were validated with the experimental data reported by Somers in NREL. Therefore, computational results for modified airfoil with sinusoidal wavy leading edge were presented. The results revealed that, at low angles of attacks before the stall region, lift coefficient decreases slightly rather than baseline model. However, the modified airfoil has a smooth stall trend while baseline airfoil lift coefficient decreases sharply due to the separation which occurred on suction side. According to the flow physics over the airfoils, leading edge bumps act as vortex generator so vortices containing high level of momentum make the flow remain attached to the surface of the airfoil at high angle of attack and prevent it from having a deep stall.

scholarly journals Active Control of a Stalled Airfoil Through Steady or Unsteady Actuation Jets

2015 ◽  
Vol 137 (9) ◽  
Author(s):  
V. G. Chapin ◽  
E. Benard
Keyword(s):  
Active Control ◽  
Computational Study ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Position Angle ◽  
Navier Stokes ◽  
Suction Surface ◽  
Parametric Studies ◽  
Naca 0012

The active control of the leading-edge (LE) separation on the suction surface of a stalled airfoil (NACA 0012) at a Reynolds number of 106 based on the chord length is investigated through a computational study. The actuator is a steady or unsteady jet located on the suction surface of the airfoil. Unsteady Reynolds-Averaged Navier–Stokes (URANS) equations are solved on hybrid meshes with the Spalart–Allmaras turbulence model. Simulations are used to characterize the effects of the steady and unsteady actuation on the separated flows for a large range of angle of attack (0 < α < 28 deg). Parametric studies are carried out in the actuator design-space to investigate the control effectiveness and robustness. An optimal actuator position, angle, and frequency for the stalled angle of attack α = 19 deg are found. A significant increase of the lift coefficient is obtained (+ 84% with respect to the uncontrolled reference flow), and the stall is delayed from angle of attack of 18 deg to more than 25 deg. The physical nonlinear coupling between the actuator position, velocity angle, and frequency is investigated. The critical influence of the actuator location relative to the separation location is emphasized.

Effect of amplitude and mean angle-of-attack on the boundary layer of an oscillating aerofoil

2008 ◽  
Vol 112 (1138) ◽  
pp. 705-713 ◽  
Author(s):  
M. R. Soltani ◽  
A. Bakhshalipour
Keyword(s):  
Boundary Layer ◽  
Transition Point ◽  
Angle Of Attack ◽  
Wind Tunnel Test ◽  
Leading Edge ◽  
Stream Velocity ◽  
Velocity Amplitude ◽  
Surface Pressure Distribution ◽  
Reduced Frequency ◽  
Hot Film

Abstract Extensive experiments were conducted to study the effect of various parameters on the surface pressure distribution and transition point of an aerofoil section used in a wind turbine blade. In this paper details of the variation of transition point on the aforementioned aerofoil are presented. The aerofoil spanned the wind-tunnel test section and was oscillated sinusoidally in pitch about the quarter chord. The imposed variables of the experiments were free stream velocity, amplitude of motion, mean angle-of-attack, and oscillation frequency. The spatial-temporal progressions of the leading-edge transition point and the state of the unsteady boundary-layer were measured using eight closely-spaced, hot-film sensors (HFS). The measurements show that: (i) Reduced frequency has a pronounced effect on the variations of the transition point. (ii) There exists a hysteresis loop in the dynamic transition location and its shape varies with the reduced frequency and mean angle-of-attack.

Lift Augmentation between Passive and Active Vortex Generator

2016 ◽  
Vol 851 ◽  
pp. 532-537
Author(s):  
Nur Faraihan Zulkefli ◽  
Zulhilmy Sahwee ◽  
Nurhayati Mohd Nur ◽  
Muhamad Nor Ashraf Mohd Fazil ◽  
Muaz Mohd Shukri
Keyword(s):  
Reynolds Number ◽  
Lift Coefficient ◽  
Vortex Generator ◽  
Leading Edge ◽  
Plasma Actuator ◽  
Triangular Shape ◽  
Subsonic Wind Tunnel ◽  
Different Types ◽  
Drag Ratio ◽  
Lift To Drag Ratio

This study was conducted to investigate the performance of passive and active vortex generator on the wing’s flap. The triangular shape of passive vortex generator (VG) was developed and attached on the wing’s flap leading edge while the plasma actuator performed as active vortex generator. The test was carried out experimentally using subsonic wind tunnel with 300 angles extended flap. Three different types of turbulent flow; with Reynolds number 1.5 x105, 2.0 x105, and 2.6x105 were used to study the aerodynamics forces of airfoil with plasma actuator OFF. All Reynolds number used were below 1x106. The result indicated that airfoil with plasma actuator produced higher lift coefficient 12% and lift-to-drag ratio 5% compared to airfoil with passive vortex generator. The overall result showed that airfoil with plasma actuator produced better lift forces compared to passive vortex generator.

Experimental Aerodynamic Characteristics of a Compound Wing in Ground Effect

2014 ◽  
Vol 136 (5) ◽  
Author(s):  
Saeed Jamei ◽  
Adi Maimun Abdul Malek ◽  
Shuhaimi Mansor ◽  
Nor Azwadi Che Sidik ◽  
Agoes Priyanto
Keyword(s):  
Reynolds Number ◽  
Drag Coefficient ◽  
Center Of Pressure ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Ground Effect ◽  
Reynolds Numbers ◽  
Aerodynamic Characteristics ◽  
Ground Clearance

Wing configuration is a parameter that affects the performance of wing-in-ground effect (WIG) craft. In this study, the aerodynamic characteristics of a new compound wing were investigated during ground effect. The compound wing was divided into three parts with a rectangular wing in the middle and two reverse taper wings with anhedral angle at the sides. The sectional profile of the wing model is NACA6409. The experiments on the compound wing and the rectangular wing were carried to examine different ground clearances, angles of attack, and Reynolds numbers. The aerodynamic coefficients of the compound wing were compared with those of the rectangular wing, which had an acceptable increase in its lift coefficient at small ground clearances, and its drag coefficient decreased compared to rectangular wing at a wide range of ground clearances, angles of attack, and Reynolds numbers. Furthermore, the lift to drag ratio of the compound wing improved considerably at small ground clearances. However, this improvement decreased at higher ground clearance. The drag polar of the compound wing showed the increment of lift coefficient versus drag coefficient was higher especially at small ground clearances. The Reynolds number had a gradual effect on lift and drag coefficients and also lift to drag of both wings. Generally, the nose down pitching moment of the compound wing was found smaller, but it was greater at high angle of attack and Reynolds number for all ground clearance. The center of pressure was closer to the leading edge of the wing in contrast to the rectangular wing. However, the center of pressure of the compound wing was later to the leading edge at high ground clearance, angle of attack, and Reynolds number.

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