Turbulent Flow Over a NACA 4412 Airfoil at Angle of Attack 15 Degree

2003 ◽  
Author(s):  
O. O. Badran ◽  
H. H. Bruun
Keyword(s):  
Reverse Flow ◽  
Separation Bubble ◽  
Angle Of Attack ◽  
Flow Region ◽  
Shear Stresses ◽  
Separation Flow ◽  
Reynolds Stresses ◽  
Mean Flow ◽  
Trailing Edge ◽  
Short Period

This paper presents the measured mean flow and Reynolds stresses results, obtained on the center-line plane of the airfoil, covering the boundary layers over the upper surface, the potential flow region and the wake downstream of the trailing edge, at αa = 15°. The flying X-hot-wire probe was used to measure the U and V components of the flow field over the airfoil. An improved understanding of the physical characteristics of separation on the airfoil sections and in the region of the trailing edge is of direct value for the improvement of high lift wings for aircraft. From the study of the separation flow at angle of attack αa = 15°, the following can be concluded: (1) An intermittent reverse flow region occurred near the trailing edge of the airfoil. A separation bubble occurred for a short period of time and was then swept away with the stream wise flow. (2) The angle of attack αa = 15° corresponds to the position of maximum lift for a NACA 4412 airfoil section. (3) It is found that values of the Reynolds normal and shear stresses move away from the surface with downstream distance, and (4) In the wake region, relatively large values of Reynolds stresses occurred, which were related to the vertical oscillation in the lower wake.

Experimental Investigation of Turbulent Flow With Separation Over a NACA 4412 Airfoil at Angle of Attack 20 Degree

2002 ◽  
Author(s):  
Omar O. Badran ◽  
Hans H. Bruun
Keyword(s):  
Separation Bubble ◽  
Angle Of Attack ◽  
Shear Stresses ◽  
Turbulence Structure ◽  
Separation Flow ◽  
Reynolds Stresses ◽  
Mean Flow ◽  
Trailing Edge ◽  
Mean Velocity ◽  
Airfoil Surface

This paper presents the measured mean flow and Reynolds stresses results, obtained on the center-line plane of the airfoil, covering the boundary layers over the upper surface, the potential flow region and the wake downstream of the trailing edge, at αa = 20°. The flying X-hot-wire probe was used to measure mean velocity and turbulence structure over the airfoil. An improved understanding of the physical characteristics of separation on the airfoil sections and in the region of the trailing edge is of direct value for the improvement of high lift wings for aircraft. From the study of the separation flow at angle of attack αa = 20°, the following can be concluded: A stable separation bubble has developed near the trailing edge of the airfoil, covering around 0.6c of the airfoil surface. Also it is found that values of the Reynolds normal and shear stresses move away from the surface with downstream distance, showing turbulence diffusion to be more evident in this flow. In the wake region, relatively large values of Reynolds stresses occurred, which were related to the vertical oscillations in the upper wake.

scholarly journals Numerical investigation of the effect of airfoil thickness on onset of dynamic stall

2019 ◽  
Vol 870 ◽  
pp. 870-900 ◽  
Author(s):  
Anupam Sharma ◽  
Miguel Visbal
Keyword(s):  
Reverse Flow ◽  
Separation Bubble ◽  
Time Integration ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Flow Region ◽  
Dynamic Stall ◽  
Implicit Time ◽  
Gradual Change ◽  
Dynamic Simulations

Effect of airfoil thickness on onset of dynamic stall is investigated using large eddy simulations at chord-based Reynolds number of 200 000. Four symmetric NACA airfoils of thickness-to-chord ratios of 9 %, 12 %, 15 % and 18 % are studied. The three-dimensional Navier–Stokes solver, FDL3DI is used with a sixth-order compact finite difference scheme for spatial discretization, second-order implicit time integration and discriminating filters to remove unresolved wavenumbers. A constant-rate pitch-up manoeuver is studied with the pitching axis located at the airfoil quarter chord. Simulations are performed in two steps. In the first step, the airfoil is kept static at a prescribed angle of attack ($=4^{\circ }$). In the second step, a ramp function is used to smoothly increase the pitch rate from zero to the selected value and then the pitch rate is held constant until the angle of attack goes past the lift-stall point. The solver is verified against experiments for flow over a static NACA 0012 airfoil. Static simulation results of all airfoil geometries are also compared against XFOIL predictions with a generally favourable agreement. FDL3DI predicts two-stage transition for thin airfoils (9 % and 12 %), which is not observed in the XFOIL results. The dynamic simulations show that the onset of dynamic stall is marked by the bursting of the laminar separation bubble (LSB) in all the cases. However, for the thickest airfoil tested, the reverse flow region spreads over most of the airfoil and reaches the LSB location immediately before the LSB bursts and dynamic stall begins, suggesting that the stall could be triggered by the separated turbulent boundary layer. The results suggest that the boundary between different classifications of dynamic stall, particularly leading edge stall versus trailing edge stall, is blurred. The dynamic-stall onset mechanism changes gradually from one to the other with a gradual change in some parameters, in this case, airfoil thickness.

Influence of Oscillating Boundary Layer Suction on Aerodynamic Performance in Highly-Loaded Compressor Cascades

2018 ◽  
Author(s):  
Xu Hao ◽  
Liu Bao ◽  
Cai Le ◽  
Zhou Xun ◽  
Wang Songtao ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Vortex Shedding ◽  
Large Scale ◽  
Reverse Flow ◽  
Separation Bubble ◽  
Flow Region ◽  
Flow Fields ◽  
Separation Flow ◽  
Oscillating Boundary ◽  
Pod Analysis

Vortex structures of the separation flow fields in compressor cascades controlled by the boundary layer oscillating suction (BLOS) are numerically investigated. The proper orthogonal decomposition (POD) method is adopted to present the variation of characteristics owned by large-scale vortices. It is found that unsteady perturbation re-organizes the aspirated flow fields and, if in a proper situation, reduces the loss furthermore. Through POD analysis, variations of vortical structures are described. The results turn out that the periodic perturbation leads to a vortex shedding process with the same frequency as the excitation. The reason of loss reduction could be summarized by actuated vortices enhancing the momentum of the stagnated fluid in the reverse flow region as well as decreasing the frequencies of vortex shedding. Finally, 3-D numerical results turn out that the oscillation can transform the stable corner separation bubble to vortex rings shedding downstream and hence improve cascade performance.

Secondary Flow Measurements in a Turbine Cascade With High Inlet Turbulence

1992 ◽  
Vol 114 (1) ◽  
pp. 173-183 ◽  
Author(s):  
D. G. Gregory-Smith ◽  
J. G. E. Cleak
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Flow Region ◽  
Shear Stresses ◽  
Reynolds Stresses ◽  
Mean Flow ◽  
Flow Measurements ◽  
Frequency Spectra ◽  
The Mean ◽  
End Wall

Measurements of the mean and turbulent flow field have been made in a cascade of high turning turbine rotor blades. The inlet turbulence was raised to 5 percent by a grid placed upstream of the cascade, and the secondary flow region was traversed within and downstream of the blades using a five-hole probe and crossed hot wires. Flow very close to the end wall was measured using a single wire placed at several orientations. Some frequency spectra of the turbulence were also obtained. The results show that the mean flow field is not affected greatly by the high inlet turbulence. The Reynolds stresses were found to be very high, particularly in the loss core. Assessment of the contributions to production of turbulence by the Reynolds stresses shows that the normal stresses have significant effects, as do the shear stresses. The calculation of eddy viscosity from two independent shear stresses shows it to be fairly isotropic in the loss core. Within the blade passage, the flow close to the end wall is highly skewed and exhibits generally high turbulence. The frequency spectra show no significant resonant peaks, except for one at very low frequency, attributable to an acoustic resonance.

Secondary Flow Measurements in a Turbine Cascade With High Inlet Turbulence

1990 ◽  
Author(s):  
D. G. Gregory-Smith ◽  
J. G. E. Cleak
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Flow Region ◽  
Shear Stresses ◽  
Reynolds Stresses ◽  
Mean Flow ◽  
Flow Measurements ◽  
Frequency Spectra ◽  
The Mean ◽  
End Wall

Measurements of the mean and turbulent flow field have been made in a cascade of high turning turbine rotor blades. The inlet turbulence was raised to 5% by a grid placed upstream of the cascade, and the secondary flow region was traversed within and downstream of the blades using a 5 hole probe and crossed hot-wires. Flow very close to the end wall was measured using a single wire placed at several orientations. Some frequency spectra of the turbulence were also obtained. The results shows that the mean flow field is not affected greatly by the high inlet turbulence. The Reynolds stresses were found to be very high, particularly in the loss core. Assessment of the contributions to production of turbulence by the Reynolds stresses show that the normal stresses have significant affects as well as the shear stresses. The calculation of eddy viscosity from two independent shear stresses show it to be fairly isotropic in the loss core. Within the blade passage, the flow close to the end wall is highly skewed and exhibits generally high turbulence. The frequency spectra show no significant resonant peaks, except for one at very low frequency, attributable to an acoustic resonance.

scholarly journals Numerical Analysis of an Instrumented Turbine Blade Cascade

2018 ◽  
Author(s):  
Bryn N. Ubald ◽  
Jiahuan Cui ◽  
Rob Watson ◽  
Paul G. Tucker ◽  
Shahrokh Shahpar
Keyword(s):  
Numerical Analysis ◽  
Measurement Accuracy ◽  
Free Stream ◽  
Separation Bubble ◽  
Leading Edge ◽  
Pressure Probe ◽  
Mean Flow ◽  
Trailing Edge ◽  
Free Stream Turbulence ◽  
Jet Trajectory

The measurement accuracy of the temperature/pressure probe mounted at the leading edge of a turbine/compressor blade is crucial for estimating the fuel consumption of a turbo-fan engine. Apart from the measurement error itself, the probe also introduces extra losses. This again would compromise the measurement accuracy of the overall engine efficiency. This paper utilizes high-fidelity numerical analysis to understand the complex flow around the probe and quantify the loss sources due to the interaction between the blade and its instrumentation. With the inclusion of leading edge probes, three dimensional flow phenomena develop, with some flow features acting in a similar manner to a jet in cross flow. The separated flow formed at the leading edge of the probe blocks a large area of the probe bleed-hole, which is one of the reasons why the probe accuracy can be sensitive to Mach and Reynolds numbers. The addition of 4% free stream turbulence is shown to have a marginal impact on the jet trajectory originated from the probe bleedhole. However, a slight reduction is observed in the size of the separation bubble formed at the leading edge of the probe, preceding the two bleedhole exits. The free stream turbulence also has a significant impact on the size of the separation bubble near the trailing edge of the blade. With the addition of the free stream turbulence, the loss observed within the trailing edge wake is reduced. More than 50% of the losses at the cascade exit are generated by the leading edge probe. A breakdown of the dissipation terms from the mean flow kinetic energy equation demonstrates that the Reynolds stresses are the key terms in dissipating the counter rotating vortex pairs with the viscous stresses responsible for the boundary layer.

scholarly journals Control of flow separation over a circular cylinder using synthetic jet

2021 ◽  
Vol 2119 (1) ◽  
pp. 012025
Author(s):  
A. S. Lebedev ◽  
M. I. Sorokin ◽  
D. M. Markovich
Keyword(s):  
Circular Cylinder ◽  
Flow Control ◽  
Flow Separation ◽  
Gas Turbines ◽  
Acoustic Noise ◽  
Reverse Flow ◽  
Flow Region ◽  
Synthetic Jet ◽  
Separation Flow ◽  
Longitudinal Length

Abstract The development of methods of active separation flow control is of great applied importance for many technical and engineering applications. Understanding the conditions for the flow separation from the surface of a bluff body is essential for the design of aircrafts, cars, hydro and gas turbines, bridges and buildings. Drag, acoustic noise, vibrations and active flow mixing depend drastically on the parameters of the vortex separation process. We investigated the possibility of reducing the longitudinal length of a reverse-flow region using the method of «synthetic jet» active separation flow control. The experiment was carried out on a compact straight-through wind channel with a 1-m long test section of a cross-section of 125x125 mm. The jet was placed at the rear stagnation point of a circular cylinder. The Reynolds number, based on the cylinder diameter and the free-stream velocity, was 5000 and the von Kármán street shedding frequency without the synthetic jet was equal to 64.8 Hz. For the first time, for such a set of parameters, we applied high speed PIV to demonstrate that the injection of the synthetic jet into the cylinder wake region leads to a significant reduction in the longitudinal length of the reverse-flow region.

Large-eddy simulation of turbulent flow over a parametric set of bumps

2019 ◽  
Vol 866 ◽  
pp. 503-525 ◽  
Author(s):  
Racheet Matai ◽  
Paul Durbin
Keyword(s):  
Large Eddy Simulation ◽  
Turbulent Flow ◽  
Pressure Gradient ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Windward Side ◽  
Reynolds Stresses ◽  
Mean Flow ◽  
Eddy Simulation ◽  
Large Eddy

Turbulent flow over a series of increasingly high, two-dimensional bumps is studied by well-resolved large-eddy simulation. The mean flow and Reynolds stresses for the lowest bump are in good agreement with experimental data. The flow encounters a favourable pressure gradient over the windward side of the bump, but does not relaminarize, as is evident from near-wall fluctuations. A patch of high turbulent kinetic energy forms in the lee of the bump and extends into the wake. It originates near the surface, before flow separation, and has a significant influence on flow development. The highest bumps create a small separation bubble, whereas flow over the lowest bump does not separate. The log law is absent over the entire bump, evidencing strong disequilibrium. This dataset was created for data-driven modelling. An optimization method is used to extract fields of variables that are used in turbulence closure models. From this, it is shown how these models fail to correctly predict the behaviour of these variables near to the surface. The discrepancies extend further away from the wall in the adverse pressure gradient and recovery regions than in the favourable pressure gradient region.

Calculation of turbulence-driven secondary motion in non-circular ducts

1984 ◽  
Vol 140 ◽  
pp. 189-222 ◽  
Author(s):  
A. O. Demuren ◽  
W. Rodi
Keyword(s):  
Eddy Viscosity ◽  
Three Dimensional ◽  
Numerical Procedure ◽  
Momentum Equation ◽  
Shear Stresses ◽  
Reynolds Stresses ◽  
Calculation Methods ◽  
Mean Flow ◽  
Square Duct ◽  
Secondary Motion

Experiments on and calculation methods for flow in straight non-circular ducts involving turbulence-driven secondary motion are reviewed. The origin of the secondary motion and the shortcomings of existing calculation methods are discussed. A more refined model is introduced, in which algebraic expressions are derived for the Reynolds stresses in the momentum equations for the secondary motion by simplifying the modelled Reynolds-stress equations of Launder, Reece & Rodi (1975), while a simple eddy-viscosity model is used for the shear stresses in the axial momentum equation. The kinetic energy k and the dissipation rate ε of the turbulent motion which appear in the algebraic and the eddy-viscosity expressions are determined from transport equations. The resulting set of equations is solved with a forward-marching numerical procedure for three-dimensional shear layers. The model, as well as a version proposed by Naot & Rodi (1982), is tested by application to developing flow in a square duct and to developed flow in a partially roughened rectangular duct investigated experimentally by Hinze (1973). In both cases, the main features of the mean-flow and the turbulence quantities are simulated realistically by both models, but the present model underpredicts the secondary velocity while the Naot-Rodi model tends to overpredict it.

Dielectric Barrier Discharge Actuators to Control Goertler Vortices on a Swept Wing

2016 ◽  
Author(s):  
Vaibhav Kumar ◽  
Nikolaus Thorell ◽  
Dhwanil Shukla ◽  
Narayanan Komerath
Keyword(s):  
Dielectric Barrier Discharge ◽  
Scaling Laws ◽  
Barrier Discharge ◽  
Delta Wing ◽  
Reverse Flow ◽  
Laser Sheet ◽  
Angle Of Attack ◽  
Flow Region ◽  
Peak Frequency ◽  
Dielectric Barrier

Narrowband excitation of fin buffeting is known to exist on several modern aircraft configurations at high angles of attack. For a fixed angle of attack and model geometry, narrowband peak frequency is a linear function of freestream speed. Under these conditions, counter-rotating vortex pairs conforming to the Goertler vortex mechanism are known to develop and amplify in the flowfield over the wings. This phenomenon is explored for relevance to reverse flow over rotor blades at high speeds as well. A 42-degree delta wing with rounded leading edges is used in a low-speed wind tunnel to confirm the phenomenon. The presence of a non-zero yaw angle can increase the strength of the Goertler vortices and also change the location of maximum intensity. Small fences on the surface have been shown to eliminate these narrow-band fluctuations. Dielectric Barrier Discharge plasma actuators offer a possible means to eliminate the narrowband excitation without obtrusive surface fences. An array of such actuators is used to generate counter-rotating vortices. Incense smoke entrained into the flow is illuminated with a laser sheet from a laser pointer. Video images are used to capture velocity in the potential flow region around the vortices. The induced velocity is used to calculate vortex strength. Scaling laws are used to estimate the frequency of the actuators, as well as the magnitude of the velocity. The scaling estimation shows that a plasma actuator is viable for model-scale configurations. Continuing experiments for the final paper plan to apply the actuator under delta wing high angle of attack operation.

Sign in / Sign up

Export Citation Format

Share Document