Steady longitudinal vortices in supersonic turbulent separated flows

2011 ◽  
Vol 672 ◽  
pp. 451-476 ◽  
Author(s):  
ERICH SCHÜLEIN ◽  
VICTOR M. TROFIMOV
Keyword(s):  
High Speed ◽  
Large Scale ◽  
Roughness Element ◽  
Vortex Generator ◽  
Vortex Pair ◽  
Leading Edge ◽  
Separated Flows ◽  
Vortex Generators ◽  
Longitudinal Vortices ◽  
Oil Flow

Large-scale longitudinal vortices in high-speed turbulent separated flows caused by relatively small irregularities at the model leading edges or at the model surfaces are investigated in this paper. Oil-flow visualization and infrared thermography techniques were applied in the wind tunnel tests at Mach numbers 3 and 5 to investigate the nominally 2-D ramp flow at deflection angles of 20°, 25° and 30°. The surface contour anomalies have been artificially simulated by very thin strips (vortex generators) of different shapes and thicknesses attached to the model surface. It is shown that the introduced streamwise vortical disturbances survive over very large downstream distances of the order of 104 vortex-generator heights in turbulent supersonic flows without pressure gradients. It is demonstrated that each vortex pair induced in the reattachment region of the ramp is definitely a child of a vortex pair, which was generated originally, for instance, by the small roughness element near the leading edge. The dependence of the spacing and intensity of the observed longitudinal vortices on the introduced disturbances (thickness and spanwise size of vortex generators) and on the flow parameters (Reynolds numbers, boundary-layer thickness, compression corner angles, etc.) has been shown experimentally.

Comparative effects of vortex generators on Ahmed’s squareback and minivan car models

2017 ◽  
Vol 231 (9) ◽  
pp. 1287-1293 ◽  
Author(s):  
Antoine Evrard ◽  
Olivier Cadot ◽  
Christophe Sicot ◽  
Vincent Herbert ◽  
Denis Ricot ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Large Scale ◽  
Vortex Generator ◽  
Vortex Generators ◽  
Longitudinal Vortices ◽  
Car Model ◽  
Bubble Length ◽  
Base Drag ◽  
Scale Roughness ◽  
Ahmed Body

This work aims to evaluate the base suction and drag modifications caused by a boundary layer manipulation due to large scale roughness prior its salient separation. A real car model, a Peugeot 208, and a squareback Ahmed body are both tested and compared in a scale 1 wind tunnel at 120 km/h with road effect and rotating wheels. The roughnesses are vortex generators placed in the boundary layer that develops on the roof of the model. They produce longitudinal vortices in the free shear. Two types of vortex generators are used, wall mounted cylinders for weak disturbances and inclined blades for stronger disturbances. It is found that whatever the vehicle is, the drag is always increased. For the squareback Ahmed body, the base suction is decreased with similar magnitudes for both vortex generators showing a beneficial effect of the vortex generator on the base drag. On the contrary, the base suction is always increased for the real car whatever the vortex generators used. In that case the effects of magnitude depends on the vortex generator types. While the cylinders degrade slightly the base suction with almost no modification in the wake, the blades are able to reduce considerably the bubble length causing a huge increase in drag, lift and base suction.

Experimental investigation of plasma vortex generator in flow control

2018 ◽  
Vol 92 (3) ◽  
pp. 376-385
Author(s):  
Alireza Ghayour ◽  
Mahmoud Mani
Keyword(s):  
Vortex Generator ◽  
Leading Edge ◽  
Streamwise Vortex ◽  
Flow Visualisation ◽  
Content Type ◽  
Oil Flow ◽  
Order Of Magnitude ◽  
Oil Flow Visualization ◽  
Unsteady Operation ◽  
Naca 0012

Purpose The purpose of this paper is to compare the effects of two different configurations of plasma streamwise vortex generators (PSVG), including comb-type and mesh-type in controlling flow. This is demonstrated on the NACA 0012 airfoil. Design/methodology/approach The investigations have been done experimentally at the various electric and aerodynamic conditions. The surface oil flow visualization method has been used to the better understanding of the flow physics and the interaction of the oncoming flow passing over the airfoil and the vortex generated by comb-type PSVG. Findings This paper demonstrates the potential capabilities of the mesh-type and comb-type PSVGs in controlling flow in unsteady operation. It was found that the vortex generated by the mesh-type PSVG in unsteady operation was an order of magnitude stronger than comb-type PSVG. The flow visualisation technic proved that only a part of the plasma actuator is effective in the condition that the actuator is installed only on a portion of the upper surface of the airfoil. Originality/value This paper experimentally confirms the capabilities of the mesh-type PSVG unsteady operation in compare with comb-type PSVG in controlling flow, whereby recommends using mesh-type PSVG in the leading edge in front of comb-type PSVG on the entire wingspan to prevent the stall.

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.

Visualizations of Flow Structures in the Rotor Passage of an Axial Compressor at the Onset of Stall

2017 ◽  
Vol 139 (4) ◽  
Author(s):  
Huang Chen ◽  
Yuanchao Li ◽  
David Tan ◽  
Joseph Katz
Keyword(s):  
High Speed ◽  
Large Scale ◽  
Incidence Angle ◽  
Axial Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Stereoscopic Particle Image Velocimetry ◽  
High Speed Imaging ◽  
Vortical Structures ◽  
Flow Phenomena

Experiments preformed in the JHU refractive index matched facility examine flow phenomena developing in the rotor passage of an axial compressor at the onset of stall. High-speed imaging of cavitation performed at low pressures qualitatively visualizes vortical structures. Stereoscopic particle image velocimetry (SPIV) measurements provide detailed snapshots and ensemble statistics of the flow in a series of meridional planes. At prestall condition, the tip leakage vortex (TLV) breaks up into widely distributed intermittent vortical structures shortly after rollup. The most prominent instability involves periodic formation of large-scale backflow vortices (BFVs) that extend diagonally upstream, from the suction side (SS) of one blade at midchord to the pressure side (PS) near the leading edge of the next blade. The 3D vorticity distributions obtained from data recorded in closely spaced planes show that the BFVs originate form at the transition between the high circumferential velocity region below the TLV center and the main passage flow radially inward from it. When the BFVs penetrate to the next passage across the tip gap or by circumventing the leading edge, they trigger a similar phenomenon there, sustaining the process. Further reduction in flow rate into the stall range increases the number and size of the backflow vortices, and they regularly propagate upstream of the leading edge of the next blade, where they increase the incidence angle in the tip corner. As this process proliferates circumferentially, the BFVs rotate with the blades, indicating that there is very little through flow across the tip region.

scholarly journals Shock-induced separated flows on the lee surface of delta wings

1987 ◽  
Vol 91 (903) ◽  
pp. 128-141 ◽  
Author(s):  
S. N. Seshadri ◽  
K. Y. Narayan
Keyword(s):  
Pressure Rise ◽  
Leading Edge ◽  
Experimental Results ◽  
Separated Flows ◽  
Delta Wings ◽  
Number Range ◽  
Pressure Distributions ◽  
Boundary Layer Interaction ◽  
Oil Flow ◽  
Attached Flow

Experiments were conducted to study shock-induced separated flows on the lee surface of delta wings with sharp leading edge at supersonic speeds. Two sets of delta wings of different thickness (10° and 25° normal angle), each with leading edge sweep angles varying from 45° to 70°, were tested. The measurements, carried out in a Mach number range from 1.4 to 3.0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wings only). Using the test results, some features of shock-induced separated flows, including in particular the boundary between this type of flow and fully attached flow, have been determined. The experimental results indicate that this boundary does not seem to show any significant dependence on wing thickness within the limit of thicknesses tested. It is shown that this boundary can be predicted for thin delta wings using a well known criterion for incipient separation in a glancing shock wave boundary layer interaction, namely that a pressure rise of 1.5 is required across the shock. Comparison of the predicted boundary with experimental results (from oil flow visualisations) shows good agreement.

Direct numerical simulation of high-speed transition due to an isolated roughness element

2014 ◽  
Vol 748 ◽  
pp. 848-878 ◽  
Author(s):  
Pramod K. Subbareddy ◽  
Matthew D. Bartkowicz ◽  
Graham V. Candler
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
High Speed ◽  
Large Scale ◽  
Stokes Equations ◽  
Roughness Element ◽  
Dynamic Mode Decomposition ◽  
Large Reynolds Number ◽  
Dynamic Mode ◽  
Low Dissipation

AbstractWe study the transition of a Mach 6 laminar boundary layer due to an isolated cylindrical roughness element using large-scale direct numerical simulations (DNS). Three flow conditions, corresponding to experiments conducted at the Purdue Mach 6 quiet wind tunnel are simulated. Solutions are obtained using a high-order, low-dissipation scheme for the convection terms in the Navier–Stokes equations. The lowest Reynolds number ($Re$) case is steady, whereas the two higher $Re$ cases break down to a quasi-turbulent state. Statistics from the highest $Re$ case show the presence of a wedge of fully developed turbulent flow towards the end of the domain. The simulations do not employ forcing of any kind, apart from the roughness element itself, and the results suggest a self-sustaining mechanism that causes the flow to transition at a sufficiently large Reynolds number. Statistics, including spectra, are compared with available experimental data. Visualizations of the flow explore the dominant and dynamically significant flow structures: the upstream shock system, the horseshoe vortices formed in the upstream separated boundary layer and the shear layer that separates from the top and sides of the cylindrical roughness element. Streamwise and spanwise planes of data were used to perform a dynamic mode decomposition (DMD) (Rowley et al., J. Fluid Mech., vol. 641, 2009, pp. 115–127; Schmid, J. Fluid Mech., vol. 656, 2010, pp. 5–28).

CFD Analysis to Investigate the Effect of Vortex Generators on a Transonic Axial Flow Compressor Stage

2015 ◽  
Author(s):  
Avinash Kumar Rajendran ◽  
M. T. Shobhavathy ◽  
R. Ajith Kumar
Keyword(s):  
Pressure Ratio ◽  
Cross Flow ◽  
Axial Flow ◽  
Vortex Generator ◽  
Leading Edge ◽  
Vortex Generators ◽  
Stall Margin ◽  
Compressor Rotor ◽  
Flow Compressor ◽  
End Wall

The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. Computational Fluid Dynamics (CFD) has been extensively used to analyze the flow through rotating machineries, in general and through axial compressors, in particular. The present work is focused on the studying the effects of Vortex Generator (VG) on test compressor at CSIR National Aerospace Laboratories, Bangalore, India using CFD. The compressor consists of NACA transonic rotor with 21 blades and subsonic stator with 18 vanes. The design pressure ratio is 1.35 at 12930 RPM with a mass flow rate of 22 kg/s. Three configurations of counter rotating VGs were selected for the analysis with 0.25δ, 0.5δ and δ height, where δ was equal to the physical thickness of boundary layer (8mm) at inlet to the compressor rotor [11]. The vortex generators were placed inside the casing at 18 percent of the chord ahead to the leading edge of the rotor. A total of 63 pairs of VGs were incorporated, with three pairs in one blade passage. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 3.5% from baseline at design speed. The reasons for this numerical stall margin improvement are discussed in detail.

High-speed laminar flow past a fin–body junction

2013 ◽  
Vol 737 ◽  
pp. 19-55 ◽  
Author(s):  
O. R. Tutty ◽  
G. T. Roberts ◽  
P. H. Schuricht
Keyword(s):  
Laminar Flow ◽  
Flat Plate ◽  
High Speed ◽  
Leading Edge ◽  
Limited Range ◽  
Interaction Region ◽  
Liquid Crystal Thermography ◽  
Heating Effects ◽  
Oil Flow ◽  
Free Stream Mach Number

AbstractInterference heating effects generated by a blunt fin-type protuberance on a flat plate exposed to a hypersonic flow have been investigated experimentally and numerically. Experiments and simulations were carried out at a free-stream Mach number of 6.7 under laminar flow conditions. The surface heating on the plate was measured experimentally using liquid-crystal thermography, which provided quantitative data with high spatial resolution. Complementary surface oil flow and schlieren experiments were also carried out to gain a better understanding of the interference flow field. The effects of fin leading-edge diameter on the heating distribution on the flat plate surface were explored. The results of the experiments and simulations agree well and reveal a highly complex interaction region which extends over seven diameters upstream of the fin. Within the interaction region surrounding the fin, heating enhancements up to ten times the undisturbed flat plate value were estimated from the experimental data. However, the liquid crystals have a limited range, and the numerical simulations indicated localized peak heating many times this value both on the plate and the fin itself.

The Response of Injection of Vortex Generator Jets to a Backward Facing Step Flow

2003 ◽  
Author(s):  
Koichi Yamagata ◽  
Manabu Saito ◽  
Tadashi Morioka ◽  
Shinji Honami
Keyword(s):  
Shear Layer ◽  
Large Scale ◽  
Flow Behavior ◽  
Velocity Ratio ◽  
Vortex Generator ◽  
Step Response ◽  
Backward Facing Step ◽  
Longitudinal Vortices ◽  
Step Flow ◽  
Vortex Generator Jets

In this paper, the flow behavior of a reattachment process over a backward facing step flow is reported. The reattachment process is controlled by injection of vortex generator jets. The injection of jets upstream of the step produces the co-rotating longitudinal vortices in a separating shear layer. The experiment of the step response of the injection jet is also conducted in order to investigate the evolution process of the longitudinal vortices. A large scale of primary and counter vortices are observed, when the velocity ratio of the free stream to injected jet is 6. The detailed structure of the longitudinal vortices is clarified. The remarkable effect of the vortices on the separating shear layer downstream of the step is observed.

scholarly journals Numerical Simulation of the Boundary Layer Control on the NACA 0015 Airfoil Through Vortex Generators

2020 ◽  
Author(s):  
Douglas da Silva ◽  
Vinicius Malatesta
Keyword(s):  
Boundary Layer ◽  
Vortex Generator ◽  
Vortex Generators ◽  
Boundary Layer Control ◽  
Recirculation Bubble ◽  
Longitudinal Vortices ◽  
Before And After ◽  
Drag And Lift ◽  
Naca 0015 ◽  
Layer Control

This paper studies the influence caused by a vortex generator (VG) on a wing section with NACA 0015 airfoil when this generator is located before and after a recirculation bubble caused by the boundary layer detachment. The study was numerically carried out and concentrated under conditions of flow with Rec = 2.38 × 105 and angles of attack AoA = 3 and 6, characterized by the fact that they undergo detachment of the boundary layer before and after the location of the VG, respectively. The use of the generator in AoA = 3 strongly influenced the reduction of the recirculation bubble, leading to a drag reduction of 1.43%. In AoA = 6 with a bubble recirculation, the effect was much lower, with no well-defined formation of longitudinal vortices, resulting in increased drag and lift at 0.33 and 0.35%, respectively.

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