An Experimental Study of the Laminar Flow Separation on a Low-Reynolds-Number Airfoil

2008 ◽  
Vol 130 (5) ◽  
Author(s):  
Hui Hu ◽  
Zifeng Yang
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Laminar Boundary Layer ◽  
Flow Separation ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Laminar Separation Bubble ◽  
Airfoil Surface ◽  
Laminar Separation ◽  
Turbulence Transition

An experimental study was conducted to characterize the transient behavior of laminar flow separation on a NASA low-speed GA (W)-1 airfoil at the chord Reynolds number of 70,000. In addition to measuring the surface pressure distribution around the airfoil, a high-resolution particle image velocimetry (PIV) system was used to make detailed flow field measurements to quantify the evolution of unsteady flow structures around the airfoil at various angles of attack (AOAs). The surface pressure and PIV measurements clearly revealed that the laminar boundary layer would separate from the airfoil surface, as the adverse pressure gradient over the airfoil upper surface became severe at AOA≥8.0deg. The separated laminar boundary layer was found to rapidly transit to turbulence by generating unsteady Kelvin–Helmholtz vortex structures. After turbulence transition, the separated boundary layer was found to reattach to the airfoil surface as a turbulent boundary layer when the adverse pressure gradient was adequate at AOA<12.0deg, resulting in the formation of a laminar separation bubble on the airfoil. The turbulence transition process of the separated laminar boundary layer was found to be accompanied by a significant increase of Reynolds stress in the flow field. The reattached turbulent boundary layer was much more energetic, thus more capable of advancing against an adverse pressure gradient without flow separation, compared to the laminar boundary layer upstream of the laminar separation bubble. The laminar separation bubble formed on the airfoil upper surface was found to move upstream, approaching the airfoil leading edge as the AOA increased. While the total length of the laminar separation bubble was found to be almost unchanged (∼20% of the airfoil chord length), the laminar portion of the separation bubble was found to be slightly stretched, and the turbulent portion became slightly shorter with the increasing AOA. After the formation of the separation bubble on the airfoil, the increase rate of the airfoil lift coefficient was found to considerably degrade, and the airfoil drag coefficient increased much faster with increasing AOA. The separation bubble was found to burst suddenly, causing airfoil stall, when the adverse pressure gradient became too significant at AOA>12.0deg.

On the origin of the inflectional instability of a laminar separation bubble

2009 ◽  
Vol 629 ◽  
pp. 263-298 ◽  
Author(s):  
SOURABH S. DIWAN ◽  
O. N. RAMESH
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Shear Layer ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Maximum Height ◽  
Laminar Separation Bubble ◽  
Laminar Separation ◽  
Separated Shear Layer ◽  
The Mean

This is an experimental and theoretical study of a laminar separation bubble and the associated linear stability mechanisms. Experiments were performed over a flat plate kept in a wind tunnel, with an imposed pressure gradient typical of an aerofoil that would involve a laminar separation bubble. The separation bubble was characterized by measurement of surface-pressure distribution and streamwise velocity using hot-wire anemometry. Single component hot-wire anemometry was also used for a detailed study of the transition dynamics. It was found that the so-called dead-air region in the front portion of the bubble corresponded to a region of small disturbance amplitudes, with the amplitude reaching a maximum value close to the reattachment point. An exponential growth rate of the disturbance was seen in the region upstream of the mean maximum height of the bubble, and this was indicative of a linear instability mechanism at work. An infinitesimal disturbance was impulsively introduced into the boundary layer upstream of separation location, and the wave packet was tracked (in an ensemble-averaged sense) while it was getting advected downstream. The disturbance was found to be convective in nature. Linear stability analyses (both the Orr–Sommerfeld and Rayleigh calculations) were performed for mean velocity profiles, starting from an attached adverse-pressure-gradient boundary layer all the way up to the front portion of the separation-bubble region (i.e. up to the end of the dead-air region in which linear evolution of the disturbance could be expected). The conclusion from the present work is that the primary instability mechanism in a separation bubble is inflectional in nature, and its origin can be traced back to upstream of the separation location. In other words, the inviscid inflectional instability of the separated shear layer should be logically seen as an extension of the instability of the upstream attached adverse-pressure-gradient boundary layer. This modifies the traditional view that pegs the origin of the instability in a separation bubble to the detached shear layer outside the bubble, with its associated Kelvin–Helmholtz mechanism. We contend that only when the separated shear layer has moved considerably away from the wall (and this happens near the maximum-height location of the mean bubble), a description by the Kelvin–Helmholtz instability paradigm, with its associated scaling principles, could become relevant. We also propose a new scaling for the most amplified frequency for a wall-bounded shear layer in terms of the inflection-point height and the vorticity thickness and show it to be universal.

Turbine Cascade Flow Control Using a “Wake Filling” Pulsed Plasma Actuator

2005 ◽  
Author(s):  
H. Perez-Blanco ◽  
Robert Van Dyken ◽  
Aaron Byerley ◽  
Tom McLaughlin
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Laminar Boundary Layer ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Plasma Actuator ◽  
Pulsed Plasma ◽  
Power Cost ◽  
Hot Film

Separation bubbles in high-camber blades under part-load conditions have been addressed via continuous and pulsed jets, and also via plasma actuators. Numerous passive techniques have been employed as well. In this type of blades, the laminar boundary layer cannot overcome the adverse pressure gradient arising along the suction side, resulting on a separation bubble. When separation is abated, a common explanation is that kinetic energy added to the laminar boundary layer speeds up its transition to turbulent. In the present study, a plasma actuator installed in the trailing edge (i.e. “wake filling configuration”) of a cascade blade is used to excite the flow in pulsed and continuous ways. The pulsed excitation can be directed to the frequencies of the large coherent structures (LCS) of the flow, as obtained via a hot-film anemometer, or to much higher frequencies present in the suction-side boundary layer, as given in the literature. It is found that pulsed frequencies much higher than that of LCS reduce losses and improve turning angles further than frequencies close to those of LCS. With the plasma actuator 50% on time, good loss abatement is obtained. Larger “on time” values yield improvements, but with decreasing returns. Continuous high-frequency activation results in the largest loss reduction, at increased power cost. The effectiveness of high frequencies may be due to separation abatement via boundary layer excitation into transition, or may simply be due to the creation of a favorable pressure gradient that averts separation as the actuator ejects fluid downstream. Both possibilities are discussed in light of the experimental evidence.

The Pervasive Effect of the Calmed Region

2005 ◽  
Author(s):  
R. L. Thomas ◽  
J. P. Gostelow
Keyword(s):  
Boundary Layer ◽  
Separation Bubble ◽  
Leading Edge ◽  
Adverse Pressure Gradient ◽  
Laminar Separation Bubble ◽  
Free Shear Layer ◽  
Hot Wire ◽  
Strong Suppression ◽  
Velocity Fluctuations ◽  
Laminar Separation

Experiments have been conducted relating to the interaction of imposed freestream wakes upon a flat plate laminar separation bubble under an adverse pressure gradient. Controlled wakes, representative of those seen in turbomachinery environments, were used to investigate unsteadiness effects upon a separating boundary layer that undergoes natural transition in the free shear layer under steady conditions. Hot-wire anemometry using a single hot-wire has shown leading edge boundary layer disturbances induced under each passing wake, which grow steadily via by-pass and natural transition methods into turbulent strips that convect with the flow. These disturbances are of such strength that the separated region is resisted and effectively swept away by the passing turbulence, momentarily giving rise to a wholly attached laminar boundary layer. Controlling the chord-wise proximity of neighboring wakes allowed for the investigation of the effect and extent of the calmed region behind each induced turbulent strip. Measurements have shown that a strong suppression of velocity fluctuations is seen related to the proximity of the turbulent strips. Turbulence level reductions of up to 40% have been demonstrated as wake spacing is reduced. Even for those cases where systematic wakes are sufficiently close together to prevent the development of a visible calmed region, very strong calming influences are seen in the wake induced turbulent domain that would have normally been occupied by the calmed flow.

Experimental Investigations of Laminar Separation Bubble for a Flow Past an Airfoil

2010 ◽  
Author(s):  
Deepakkumar M. Sharma ◽  
Kamal Poddar
Keyword(s):  
Boundary Layer ◽  
Flow Separation ◽  
Separation Bubble ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Boundary Layer Separation ◽  
Experimental Investigations ◽  
Stream Velocity ◽  
Laminar Separation Bubble ◽  
Laminar Separation

Wind tunnel experiments were carried out on NACA 0015 airfoil model to investigate the formation of laminar separation bubble on the upper surface of the airfoil by varying angle of attack from −5° to 25° with respect to the free stream velocity at constant Reynolds number varying from 0.2E06 to 0.6E06. Pressure signals were acquired from the pressure ports selected at the mid-span of the airfoil model along the chord. Static stall characteristics were obtained from the surface pressure distribution. The flow separation was found to be a trailing edge turbulent boundary layer separation preceded with a laminar separation bubble. Flow Visualizations were done by using Surface Oil flow Technique for qualitative analysis of the transition zone formed due to the presence of laminar separation bubble As the angle of attack is increased the separation bubble moves towards the leading edge of the airfoil and finally gets shredded or burst at a particular angle of attack resulting in leading edge turbulent flow separation which induces the static stall condition. The flow separation process is critically analyzed and the existence of laminar separation bubble is visualized and quantified with the increase in angle of attack and Re. Effect of Re and angle of attack on the various boundary layer and Separation bubble parameters are obtained and analyzed.

Boundary Layer Flashback Model for Hydrogen Flames in Confined Geometries Including the Effect of Adverse Pressure Gradient

2020 ◽  
Author(s):  
Ólafur H. Björnsson ◽  
Sikke A. Klein ◽  
Joeri Tober
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Flow Separation ◽  
Gradient Flow ◽  
Lewis Number ◽  
Adverse Pressure Gradient ◽  
Future Research ◽  
Diffuser Flow ◽  
Combustion Properties ◽  
Quenching Distance

Abstract The combustion properties of hydrogen make premixed hydrogen-air flames very prone to boundary layer flashback. This paper describes the improvement and extension of a boundary layer flashback model from Hoferichter [1] for flames confined in burner ducts. The original model did not perform well at higher preheat temperatures and overpredicted the backpressure of the flame at flashback by 4–5x. By simplifying the Lewis number dependent flame speed computation and by applying a generalized version of Stratford’s flow separation criterion [2], the prediction accuracy is improved significantly. The effect of adverse pressure gradient flow on the flashback limits in 2° and 4° diffusers is also captured adequately by coupling the model to flow simulations and taking into account the increased flow separation tendency in diffuser flow. Future research will focus on further experimental validation and direct numerical simulations to gain better insight into the role of the quenching distance and turbulence statistics.

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