Drag reduction of 3D bluff body using SDBD plasma actuators

2020 ◽  
pp. 095440702096189
Author(s):  
Mostafa Kazemi ◽  
Parisa Ghanooni ◽  
Mahmoud Mani ◽  
Mohammad Saeedi
Keyword(s):  
Drag Reduction ◽  
Drag Coefficient ◽  
Bluff Body ◽  
Separation Bubble ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Force Balance ◽  
Plasma Actuators ◽  
Laminar Separation ◽  
Trailing Vortices

In the current research, a series of different combinations of plasma SDBD actuators mounted on a simplified road vehicle have been experimentally studied to find the optimum position of the actuators for controlling the flow separation and reducing the vehicle form drag. Separation point of the flow over the rear ramp, large trailing vortices of the standard model, and laminar separation bubble (LSB) of the rear ramp leading edge are among the most significant factors to be controlled. The experiments were conducted at Reynolds numbers ranging from 0.55 × 106 to 1.11 × 106 in a subsonic wind tunnel while the pressure distribution over the model and its streamwise force balance were accurately measured. Significant drag reduction due to the use of DBD actuators was observed. As such, for the range of tested Reynolds numbers, a maximum of 25.1% of drag reduction in the vehicle drag coefficient could be achieved. The optimum combinations of activation voltages (6, 9, and 12 kV) and wave frequencies (6, 10, and 14 kHz) for plasma actuators were also found. Furthermore, it was observed that SDBD actuators mounted on the rear ramp of the model had a deeper impact on the vehicle drag coefficient compared to the other actuators.

Experimental Study of Flow Separation over NACA633018 Wing with Synthetic Jet Control at Low Reynolds Numbers

2012 ◽  
Vol 29 (1) ◽  
pp. 45-52 ◽  
Author(s):  
C.-Y. Lin ◽  
F.-B. Hsiao
Keyword(s):  
Reynolds Number ◽  
Flow Separation ◽  
Separation Bubble ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Synthetic Jet ◽  
Synthetic Jets ◽  
Laminar Separation ◽  
Low Reynolds Numbers ◽  
Jet Control

AbstractThis paper experimentally studies flow separation and aerodynamic performance of a NACA633018 wing using a series of piezoelectric-driven disks, which are located at 12% chord length from the leading edge to generate a spanwise-distributed synthetic jets to excite the passing flow. The experiment is conducted in an open-type wind tunnel with Reynolds numbers (Re) of 8 × 104 and 1.2 × 105, respectively, based on the wing chord. The oscillations of the synthetic jet actuators (SJAs) disturb the neighboring passage flow on the upper surface of the wing before the laminar separation takes place. The disturbances of energy influence the downstream development of boundary layers to eliminate or reduce the separation bubble on the upper surface of the wing. Significant lift increase and drag decrease are found at the tested Reynolds number of 8 × 104 due to the actuators excitation. Furthermore, the effect of drag also reduces dominant with increasing Reynolds number, but the increase on lift is reduced with the Reynolds number increased.

The Effect of Reynolds Number and Laminar Separation on Axial Cascade Performance

1975 ◽  
Vol 97 (2) ◽  
pp. 261-273 ◽  
Author(s):  
W. B. Roberts
Keyword(s):  
Reynolds Number ◽  
Shear Layer ◽  
Separation Bubble ◽  
Axial Compressor ◽  
Leading Edge ◽  
Practical Method ◽  
Reynolds Numbers ◽  
Compressor Cascade ◽  
Laminar Separation ◽  
Semiempirical Theory

Testing over a range of Reynolds numbers was done for three NACA 65 Profiles in cascade. The testing was carried out in the VKI C-1 Low Speed Cascade Wind Tunnel; blade chord Reynolds number was varied from 250,000 to 40,000. A semiempirical theory is developed which will predict the behavior of the shear layer across a laminar separation bubble. The method is proposed for two-dimensional incompressible flow and is applicable down to short bubble bursting. The method can be used to predict the length of the laminar bubble, the bursting Reynolds number, and the development of the shear layer through the separated region. As such it is a practical method for calculating the profile losses of axial compressor and turbine cascades in the presence of laminar separation bubbles. It can also be used to predict the abrupt leading edge stall associated with thin airfoil sections. The predictions made by the method are compared with the available experimental data. The agreement could be considered good. The method was also used to predict regions of laminar separation in converging flows through axial compressor cascades (exterior to the corner vortices) with good results. For Reynolds numbers below bursting the semiempirical theory no longer applies. For this situation the performance of an axial compressor cascade can be computed using an empirical correlation proposed by the author. Comparison of performance prediction with experiment shows satisfactory agreement. Finally, a tentative correlation, based on the NACA Diffusion Factor, is presented that allows a rapid estimation of the bursting Reynolds number of an axial compressor cascade.

Interaction of tip vortex and laminar separation bubble over wings with different aspect ratios under low Reynolds numbers

2018 ◽  
Vol 232 (22) ◽  
pp. 4019-4037 ◽  
Author(s):  
Mustafa Serdar Genç ◽  
Gökhan Özkan ◽  
Mustafa Özden ◽  
Mehmet Sadik Ki˙ri˙ş ◽  
Rahime Yi˙ldi˙z
Keyword(s):  
Separation Bubble ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Tip Vortex ◽  
Transition To Turbulence ◽  
Laminar Separation ◽  
Viscous Forces ◽  
Aspect Ratios ◽  
Oil Flow ◽  
2D Flow

In present study, aerodynamics of a NACA4412 wings with aspect ratio of 1 and 3 was considered experimentally at Reynolds numbers of 2.5 × 104, 5 × 104 and 7.5 × 104. Studies for AR = 1 wing showed that stall was delayed and extra (vortex) lift was obtained, because separation bubble got smaller in both chordwise and spanwise axes with effect of wing-tip vortices. Oil-flow experiments at higher angles of attack clarified the reason for vortex lift obtained from AR = 1 wing. However, there was an increase in drag coefficient as well as vortex lift, and stall delayed due to tip vortex. Turbulence intensity distributions pointed out location of the transition to turbulence; Reynolds stress and turbulence kinetic energy distributions indicated shear layer. Furthermore, in experiments of AR = 3 wing, the viscous forces and leading edge vortices were effective at Re = 2.5 × 104 and Re = 5 × 104, but flow over the wing at Re = 7.5 × 104 acted as a 2D flow. After α = 12°, bubble burst and stall consisted abruptly because effectiveness of 3D flow decreased over wing. Strouhal (St) numbers of vortex shedding frequencies in wake of AR = 3 wing had a certain difference from St = 0.17/sinα curve at lower angle of attack (α = 0° − 10°) due to separation bubble, but AR = 1 wings showed that St numbers were near St = 0.17/sinα curve.

scholarly journals Separation Control of NACA0015 Airfoil using Plasma Actuators

2021 ◽  
Vol 1 (1) ◽  
Author(s):  
Akira Aiura ◽  
Kentaro Kobayashi ◽  
Jun Sakakibara
Keyword(s):  
Separation Bubble ◽  
Leading Edge ◽  
Ccd Camera ◽  
Force Balance ◽  
Plasma Actuators ◽  
Large Field ◽  
Separation Control ◽  
Pressure Transducers ◽  
Periodic Disturbances ◽  
Two Component

Separation control of NACA0015 airfoil using plasma actuators was investigated. Plasma actuators in spanwise array, which consists of 21 electrodes, were located at the leading edge of the airfoil to give temporal periodic disturbances with phase variations into its boundary layer. The cord length of the airfoil was c = 100mm and corresponding Reynolds number was fixed at Re = 63,000. Non-dimensional frequency of the disturbance was chosen at F+ = 0.5 or 6. The gap between adjacent electrode was set as 1mm, and phase difference of the temporal periodic disturbances between adjacent electrode was set at φ = 0 or π. Velocity field was measured by conventional two-component PIV using a CCD camera (Imperx, B1922, 1920 x 1460pixels) and Nd-YAG laser (Quantel, Evergreen, 140mJ/pulse). Both large field of view (FOV) images capturing whole wing with surrounding flow and smaller FOV images focused on the separation bubble near leading edge were evaluated. Surface pressure was monitored by pressure transducers through pressure taps on the upper surface of airfoil. Lift and drag against the airfoil were measured using a two-component force balance.

An experimental investigation of a laminar separation bubble on the leading-edge of a modelled aerofoil for different Reynolds numbers

2015 ◽  
Vol 230 (13) ◽  
pp. 2208-2224 ◽  
Author(s):  
A Samson ◽  
S Sarkar
Keyword(s):  
Reynolds Number ◽  
Shear Layer ◽  
Separation Bubble ◽  
Experimental Studies ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Constant Thickness ◽  
Laminar Separation Bubble ◽  
Laminar Separation ◽  
Separated Shear Layer

This paper describes the dynamics of a laminar separation bubble formed on the semi-circular leading edge of constant thickness aerofoil model. Detailed experimental studies are carried out in a low-speed wind tunnel, where surface pressure and time-averaged velocity in the separated region and as well as in the downstream are presented along with flow field visualisations through PIV for various Reynolds numbers ranging from 25,000 to 75,000 (based on the leading edge diameter). The results illustrate that the separated shear layer is laminar up to 20% of separation length and then the perturbations are amplified in the second half attributing to breakdown and reattachment. The bubble length is highly susceptible to change in Reynolds number and plays an important role in outer layer activities. Further, the transition of a separated shear layer is studied through variation of intermittency factor and comparing with existing correlations available in the literature for attached flow and as well as separated flow. Transition of the separated shear layer occurs through formation of K-H rolls, where the intermittency following spot propagation theory appears valid. The predominant shedding frequency when normalised with respect to the momentum thickness at separation remains almost constant with change in Reynolds number. The relaxation is slow after reattachment and the flow takes about five bubble lengths to approach a canonical layer.

Investigation on the effect of leading edge tubercles of sweptback wing at low reynolds number

2020 ◽  
Vol 21 (6) ◽  
pp. 621
Author(s):  
Veerapathiran Thangaraj Gopinathan ◽  
John Bruce Ralphin Rose ◽  
Mohanram Surya
Keyword(s):  
Leading Edge ◽  
Reynolds Numbers ◽  
Sweep Angle ◽  
Laminar Separation ◽  
Low Reynolds Numbers ◽  
Flow Energy ◽  
Airplane Wing ◽  
Low Reynolds ◽  
Naca 0015 ◽  
Wing Performance

Aerodynamic efficiency of an airplane wing can be improved either by increasing its lift generation tendency or by reducing the drag. Recently, Bio-inspired designs have been received greater attention for the geometric modifications of airplane wings. One of the bio-inspired designs contains sinusoidal Humpback Whale (HW) tubercles, i.e., protuberances exist at the wing leading edge (LE). The tubercles have excellent flow control characteristics at low Reynolds numbers. The present work describes about the effect of tubercles on swept back wing performance at various Angle of Attack (AoA). NACA 0015 and NACA 4415 airfoils are used for swept back wing design with sweep angle about 30°. The modified wings (HUMP 0015 A, HUMP 0015 B, HUMP 4415 A, HUMP 4415 B) are designed with two amplitude to wavelength ratios (η) of 0.1 & 0.24 for the performance analysis. It is a novel effort to analyze the tubercle vortices along the span that induce additional flow energy especially, behind the tubercles peak and trough region. Subsequently, Co-efficient of Lift (CL), Co-efficient of Drag (CD) and boundary layer pressure gradients also predicted for modified and baseline (smooth LE) models in the pre & post-stall regimes. It was observed that the tubercles increase the performance of swept back wings by the enhanced CL/CD ratio in the pre-stall AoA region. Interestingly, the flow separation region behind the centerline of tubercles and formation of Laminar Separation Bubbles (LSB) were asymmetric because of the sweep.

Experimental investigations of flow over NACA airfoils 0021 and 4412 of wind turbine blades with and without Tubercles

2021 ◽  
pp. 0309524X2110071
Author(s):  
Usman Butt ◽  
Shafqat Hussain ◽  
Stephan Schacht ◽  
Uwe Ritschel
Keyword(s):  
Wind Tunnel ◽  
Drag Coefficient ◽  
Wind Turbine ◽  
Critical Region ◽  
Lift Coefficient ◽  
Turbine Blades ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Experimental Investigations ◽  
Wind Turbine Blades

Experimental investigations of wind turbine blades having NACA airfoils 0021 and 4412 with and without tubercles on the leading edge have been performed in a wind tunnel. It was found that the lift coefficient of the airfoil 0021 with tubercles was higher at Re = 1.2×105 and 1.69×105 in post critical region (at higher angle of attach) than airfoils without tubercles but this difference relatively diminished at higher Reynolds numbers and beyond indicating that there is no effect on the lift coefficients of airfoils with tubercles at higher Reynolds numbers whereas drag coefficient remains unchanged. It is noted that at Re = 1.69×105, the lift coefficient of airfoil without tubercles drops from 0.96 to 0.42 as the angle of attack increases from 15° to 20° which is about 56% and the corresponding values of lift coefficient for airfoil with tubercles are 0.86 and 0.7 at respective angles with18% drop.

Aerodynamic Drag Reduction Investigation for a Simplified Road Vehicle Using Plasma Flow Control

2016 ◽  
Author(s):  
Bahram Khalighi ◽  
Joanna Ho ◽  
John Cooney ◽  
Brian Neiswander ◽  
Thomas C. Corke ◽  
...  
Keyword(s):  
Drag Reduction ◽  
Flow Control ◽  
Plasma Flow ◽  
Aerodynamic Drag ◽  
Force Balance ◽  
Plasma Actuators ◽  
Ground Vehicles ◽  
Mean Velocity ◽  
Local Flow ◽  
Plasma Flow Control

The effect of plasma flow control on reducing aerodynamic drag for ground vehicles is investigated. The experiments were carried out for a simplified ground vehicle using single dielectric barrier discharge (SDBD) plasma actuators. The plasma actuators were designed to alter the flow structure in the wake region behind the vehicle. The Ahmed body was modified to allow eight different vehicle geometries (with backlight or slant angles of 0° and 35°). Each of these were further modified by rounding the edges with different radii. Flow visualizations such as particle streams and surface oil were used to quantify features of the local flow field. The drag on the models was measured using a force balance as well as by integrating the mean velocity profiles in the model wakes. The results indicated that flow modifications needed to be applied symmetrically (upper to lower and/or side to side). This was demonstrated with the 0° backlight angle (square-back) that had all four side-corners rounded. Plasma actuators were applied to all four of the rounded edges to enhance the ability to direct the flow into the wake. Wake measurements showed that steady actuation at a fixed actuator voltage reduced the drag by an average of 20% at the lower velocities (below 15 m/s) and by 3% at the highest velocity tested (20 m/s). Model constraints prevented increasing the plasma actuator voltage that was needed to maintain the higher drag reduction observed at the lower speeds.

Turbulent Spots in Strong Adverse Pressure Gradients: Part 2 — Spot Propagation and Spreading Rates

2001 ◽  
Author(s):  
A. D’Ovidio ◽  
J. A. Harkins ◽  
J. P. Gostelow
Keyword(s):  
Separation Bubble ◽  
Reynolds Numbers ◽  
Boundary Layer Transition ◽  
Pressure Gradients ◽  
Flat Plates ◽  
Layer Transition ◽  
Turbulent Spots ◽  
Laminar Separation ◽  
Transition Length ◽  
Spreading Rates

The study of turbulent spots in strong adverse pressure gradients is of current interest in turbomachinery research. The aim of this investigation is to use information gathered from boundary layer transition and laminar separation, in wind tunnel tests on flat plates, to predict the equivalent phenomena occurring on turbomachinery blade surfaces. In Part 1 turbulent spot behavior was documented for two Reynolds numbers, corresponding to a laminar separation bubble (LSB) and an incipient separation condition (IS). In Part 2 further results are reported characterizing typical spot propagation and spreading rates and serving to validate or modify existing correlations for predicting transition length.

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