Effects of Reynolds Number and Surface Roughness Magnitude and Location on Compressor Cascade Performance

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Seung Chul Back ◽  
Garth V. Hobson ◽  
Seung Jin Song ◽  
Knox T. Millsaps
Keyword(s):  
Surface Roughness ◽  
Reynolds Number ◽  
Leading Edge ◽  
Side Surface ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Pressure Side ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Blade Loading

An experimental investigation has been conducted to characterize the influence of Reynolds number and surface roughness magnitude and location on compressor cascade performance. Flow field surveys have been conducted in a low-speed, linear compressor cascade. Pressure, velocity, and loss have been measured via a five-hole probe, pitot probe, and pressure taps on the blades. Four different roughness magnitudes, Ra values of 0.38 μm (polished), 1.70 μm (baseline), 2.03 μm (rough 1), and 2.89 μm (rough 2), have been tested. Furthermore, various roughness locations have been examined. In addition to the as manufactured (baseline) and entirely rough blade cases, blades with roughness covering the leading edge, pressure side, and 5%, 20%, 35%, 50%, and 100% of suction side from the leading edge have been studied. All of the tests have been carried out for Reynolds numbers ranging from 300,000 to 640,000. For Reynolds numbers under 500,000, the tested roughnesses do not significantly degrade compressor blade loading or loss. However, loss and blade loading become sensitive to roughness at Reynolds numbers above 550,000. Cascade performance is more sensitive to roughness on the suction side than pressure side. Furthermore, roughness on the aft 2/3 of suction side surface has a greater influence on loss. For a given roughness location, there exists a Reynolds number at which loss begins to significantly increase. Finally, increasing the roughness area on the suction surface from the leading edge reduces the Reynolds number at which the loss begins to increase.

Effect of Surface Roughness Location and Reynolds Number on Compressor Cascade Performance

2010 ◽  
Author(s):  
Seung Chul Back ◽  
Garth V. Hobson ◽  
Seung Jin Song ◽  
Knox T. Millsaps
Keyword(s):  
Surface Roughness ◽  
Reynolds Number ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Trailing Edge ◽  
Blade Loading ◽  
Linear Compressor Cascade

An experimental investigation has been conducted to characterize the influence of surface roughness location and Reynolds number on compressor cascade performance. Flow field surveys have been conducted in a low-speed, linear compressor cascade. Pressure, velocity, and flow angles have been measured via a 5-hole probe, pitot probe, and pressure taps on the blades. In addition to the entirely smooth and entirely rough blade cases, blades with roughness covering the leading edge; pressure side; and 5%, 20%, 35%, 50%, and 100% of suction side from the leading edge have been studied. All of the tests have been done for Reynolds number ranging from 300,000 to 640,000.Cascade performance (i.e. blade loading, loss, and deviation) is more sensitive to roughness on the suction side than pressure side. Roughness near the trailing edge of suction side increases loss more than that near the leading edge. When the suction side roughness is located closer to the trailing edge, the deviation and loss increase more rapidly with Reynolds number. For a given roughness location, there exists a Reynolds number at which loss begins to visibly increase. Finally, increasing the area of rough suction surface from the leading edge reduces the Reynolds number at which the loss coefficient begins to increase.

Effect of Leading Edge Roughness and Reynolds Number on Compressor Profile Loss

2013 ◽  
Author(s):  
Ju Hyun Im ◽  
Ju Hyun Shin ◽  
Garth V. Hobson ◽  
Seung Jin Song ◽  
Knox T. Millsaps
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Separation Bubble ◽  
Leading Edge ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Laser Doppler Velocimeter ◽  
Compressor Cascade ◽  
Edge Roughness ◽  
Profile Loss

An experimental investigation has been conducted to characterize the influence of leading edge roughness and Reynolds number on compressor cascade profile loss. Tests have been conducted in a low-speed linear compressor cascade at Reynolds numbers between 210,000 and 640,000. Blade loading and loss have been measured with pressure taps and pneumatic probes. In addition, a two-component laser-doppler velocimeter (LDV) has been used to measure the boundary layer velocity profiles and turbulence levels at various chordwise locations near the blade suction surface. The “smooth” blade has a centerline-averaged roughness (Ra) of 0.62 μm. The “rough” blade is roughened by covering the leading edge of the “smooth” blade, including 2% of the pressure side and 2% of the suction side, with a 100 μm-thick tape with a roughness Ra of 4.97 μm. At Reynolds numbers ranging from 210,000 to 380,000, the leading edge roughness decreases loss slightly. At Reynolds number of 210,000, the leading edge roughness reduces the size of the suction side laminar separation bubble and turbulence level in the turbulent boundary layer after reattachment. Thus, the leading edge roughness reduces displacement and momentum thicknesses as well as profile loss at Reynolds number of 210,000. However, the same leading edge roughness increases loss significantly for Re = 450,000 ∼ 640,000. At Reynolds number of 640,000, the leading edge roughness decreases the magnitude of the favorable pressure gradient for axial chordwise locations less than 0.41 and induces turbulent separation for axial chordwise locations greater than 0.63, drastically increasing loss. Thus, roughness limited to the leading edge still has a profound effect on the compressor flow field.

Darryl E. Metzger Memorial Session Paper: The Influence of Secondary Flows Near the Endwall and Boundary Layer Disturbance on Convective Transport From a Turbine Blade

1995 ◽  
Vol 117 (4) ◽  
pp. 657-665 ◽  
Author(s):  
R. J. Goldstein ◽  
H. P. Wang ◽  
M. Y. Jabbari
Keyword(s):  
Mass Transfer ◽  
Transfer Rate ◽  
Mass Transfer Rate ◽  
Leading Edge ◽  
Flow Region ◽  
Suction Side ◽  
Convective Transport ◽  
Pressure Side ◽  
Suction Surface ◽  
Passage Vortex

A naphthalene sublimation technique is used to investigate convective transport from a simulated turbine blade in a stationary linear cascade. In some of the tests undertaken, a trip wire is stretched along the span of the blade near the leading edge. The disturbance produced by tripping the boundary layers on the blade near the leading edge causes early boundary layer transition, creates high mass transfer rate on the pressure side and in the laminar flow region on the suction side, but lowers the transfer rate in the turbulent flow region on the suction side. Comparison is made with other heat and mass transfer studies in the two-dimensional region far from the endwall and good agreement is found. Near the endwall, flow visualization indicates a strong secondary flow pattern. The impact of vortices initiated near the endwall on the laminar–turbulent transition extends three-dimensional effects to about 0.8 chord lengths on the suction side and to about 0.2 chord lengths on the pressure side away from the endwall. The effect of the passage vortex and the new vortex induced by the passage vortex on mass transfer is clearly seen and can be traced along the suction surface of the blade. Close to the endwall the highest mass transfer rate on the suction surface is not found near the leading edge. It occurs at about 27 percent of the curvilinear distance from the stagnation line to the trailing edge where a strong main flow and the secondary passage flow from the pressure side of the adjacent blade interact. The influences of some small but very intense corner vortices and the passage vortex on mass transfer are also observed on both surfaces of the blade.

scholarly journals The Influence of Secondary Flows Near the Endwall and Boundary Layer Disturbance on Convective Transport From a Turbine Blade

1994 ◽  
Author(s):  
R. J. Goldstein ◽  
H. P. Wang ◽  
M. Y. Jabbari
Keyword(s):  
Mass Transfer ◽  
Transfer Rate ◽  
Mass Transfer Rate ◽  
Leading Edge ◽  
Flow Region ◽  
Suction Side ◽  
Convective Transport ◽  
Pressure Side ◽  
Suction Surface ◽  
Passage Vortex

A naphthalene sublimation technique is used to investigate convective transport from a simulated turbine blade in a stationary linear cascade. In some of the tests undertaken a trip wire is stretched along the span of the blade near the leading edge. The disturbance produced by tripping the boundary layers on the blade near the leading edge causes early boundary layer transition, creates high mass transfer rate on the pressure side and in the laminar flow region on the suction side, but lowers the transfer rate in the turbulent flow region on the suction side. Comparison is made with other heat and mass transfer studies in the two dimensional region far from the endwall and good agreement is found. Near the endwall, flow visualization indicates a strong secondary flow pattern. The impact of vortices initiated near the endwall on the laminar-turbulent transition extends three dimensional effects to about 0.8 chord lengths on the suction side and to about 0.2 chord lengths on the pressure side away from the endwall. The effect of the passage vortex and the new vortex induced by the passage vortex on mass transfer is clearly seen and can be traced along the suction surface of the blade. Close to the endwall the highest mass transfer rate on the suction surface is not found near the leading edge. It occurs at about 27% of the curvilinear distance from the stagnation line to the trailing edge where a strong main flow and the secondary passage flow from the pressure side of the adjacent blade interact. The influences of some small but very intense corner vortices and the passage vortex on mass transfer are also observed on both surfaces of the blade.

scholarly journals Aerodynamic characteristics of owl like airfoil at low reynolds number.

2022 ◽  
pp. 26-34
Author(s):  
Ishfaq Fayaz ◽  
Syeeda Needa Fathima ◽  
Y.D. Dwivedi
Keyword(s):  
Reynolds Number ◽  
Separation Bubble ◽  
Low Reynolds Number ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Aerodynamic Characteristics ◽  
Pressure Side ◽  
Low Reynolds Numbers ◽  
Lift Curve ◽  
Low Reynolds

The computational investigation of aerodynamic characteristics and flow fields of a smooth owl-like airfoil without serrations and velvet structures.The bioinspired airfoil design is planned to serve as the main-wing for low-reynolds number aircrafts such as (MAV)micro air vechiles.The dependency of reynolds number on aerodynamics could be obtained at low reynolds numbers.The result of this experiment shows the owl-like airfoil is having high lift performance at very low speeds and in various wind conditions.One of the unique feature of owl airfoil is a separation bubble on the pressure side at low angle of attack.The separation bubble changes location from the pressure side to suction side as the AOA (angle of attack) increases. The reynolds number dependancy on the lift curve is insignificant,although there’s difference in drag curve at high angle of attacks.Eventually, we get the geometric features of the owl like airfoil to increase aerodynamic performance at low reynolds numbers.

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.

Formation and Transport of Secondary Flows Caused by Vortex-Blade Interaction in a High Pressure Turbine

2015 ◽  
Author(s):  
Zuo-Jun Wei ◽  
Wei-Yang Qiao ◽  
Ping-Ping Chen ◽  
Jian Liu
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Pressure Gradient ◽  
Suction Side ◽  
Secondary Flows ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Blade Loading ◽  
Passage Vortex

As modern turbines are designed with low aspect ratio and high blade loading, secondary flow interactions become more important. In the present work, numerical simulation is performed in a two-stage high-pressure turbine with divergent meridional passage to investigate the transport and interaction of secondary vortex from the first stage rotor within the second stage’s stator. Scale-Adaptive Simulation model coupled with Shear Stress Transport model (SAS-SST turbulence model) is used to capture the flow structures caused by the interaction in the second stator. Coupled with the passage vortex of the first rotor, the shed vortex rotates opposite in the direction and has comparable strength. As both of these vortices convect downstream to the stator bladerow, each deforms into two legs on the pressure and suction sides in the passage. In the passage due to the cross pressure gradient by blade loading, all the low-momentum fluid contained in these vortices moves towards the suction side. Besides, with the existing static pressure gradient in radial direction and vortex dynamics, the suction-side leg and the pressure-side leg move in different radial directions. The suction side leg of incoming passage vortex moves towards the endwall along the suction surface and interacts with the developing passage vortex of the second stator. The incoming shed vortex moves towards the midspan and rolls up the boundary layer fluid from suction surface. Due to the interactions between the incoming shed vortices from the hub and casing and the boundary layer of second stator, two counter-rotating vortices are formed near the midspan. Additional high loss is found there at the outlet plane, which has a comparable magnitude to the endwall secondary loss. The pressure side leg of the incoming passage vortex remains in a certain span with that of the incoming shed vortex and is not engulfed by the developing passage vortex.

Effect of Roughness and Unsteadiness on the Performance of a New Low Pressure Turbine Blade at Low Reynolds Numbers

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Francesco Montomoli ◽  
Howard Hodson ◽  
Frank Haselbach
Keyword(s):  
Reynolds Number ◽  
Wind Tunnel ◽  
Beneficial Effect ◽  
High Speed ◽  
Leading Edge ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
High Lift ◽  
Wake Passing

This paper presents a study of the performance of a high-lift profile for low pressure turbines at Reynolds numbers lower than in previous investigations. By following the results of Coull et al. (2008, “Velocity Distributions for Low Pressure Turbines,” ASME Paper No. GT2008-50589) on the design of high-lift airfoils, the profile is forward loaded. The separate and combined effects of roughness and wake passing are compared. On a front loaded blade, the effect of incidence becomes more important and the consequences in terms of cascade losses, is evaluated. The experimental investigation was carried out in the high speed wind tunnel of Whittle Laboratory, University of Cambridge. This is a closed-circuit continuous wind tunnel where the Reynolds number and Mach number can be fixed independently. The unsteadiness caused by wake passing in front of the blades is reproduced using a wake generator with rotating bars. The results confirm that the beneficial effect of unsteadiness on losses is present even at the lowest Reynolds number examined (Re3=20,000). This beneficial effect is reduced at positive incidence. With a front loaded airfoil and positive incidence, the transition occurs on the suction side close to the leading edge and this results in higher losses. This has been found valid for the entire Reynolds range investigated (20,000≤Re3≤140,000). Roughening the surface also had a beneficial effect on the losses but this effect vanishes at the lower Reynolds numbers, i.e., (Re3≤30,000), where the surface becomes hydraulically smooth. The present study suggests that a blade with as-cast surface roughness has a lower loss than a polished one.

Effects of Blade Tip Geometries on the Aerodynamic Performance of a Compressor Cascade

2013 ◽  
Author(s):  
Hongwei Ma ◽  
Jun Zhang ◽  
Wei Wei
Keyword(s):  
Pressure Rise ◽  
Suction Side ◽  
Leakage Flow ◽  
Pressure Side ◽  
Compressor Cascade ◽  
Turning Angle ◽  
Blade Loading ◽  
Blade Width ◽  
Blade Tip ◽  
Greater Curvature

This paper presents a numerical simulation of effects of blade tip geometries on the flow field of a compressor cascade. The tip geometries include flat tip (baseline), tip with cavity, tip with pressure side extension and suction side squealer tip. For the tip with cavity and pressure side extension, the mass of the leakage flow is reduced. The loss in the tip gap of the cavity tip is greater than the baseline because of the interaction of the cavity flow and the leakage flow. For the tip with pressure side extension, the loss in the gap is also greater than the baseline. The main reason is that the greater blade width makes the mixing process of the leakage flow in the gap more sufficient than the baseline. For both these two cases, the turning angle of the cascade becomes smaller and the pressure rise of the cascade is lower than the flat tip case. For the suction side squealer tip, the greater curvature of squealer increases the blade loading. The turning angle of the cascade and the pressure rise becomes greater which increases total pressure loss slightly.

The Effects of Inlet Reynolds Number, Exit Mach Number and Incidence Angle on Leading Edge Film Cooling Effectiveness of a Turbine Blade in a Linear Transonic Cascade

2015 ◽  
Author(s):  
Cong Liu ◽  
Hui-ren Zhu ◽  
Zhong-yi Fu ◽  
Run-hong Xu
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

This paper experimentally investigates the film cooling performance of a leading edge with three rows of film holes on an enlarged turbine blade in a linear cascade. The effects of blowing ratio, inlet Reynolds number, isentropic exit Mach number and off-design incidence angle (i<0°) are considered. Experiments were conducted in a short-duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and exit Mach number independently. The surface film cooling measurements were made at the midspan of the blade using thermocouples based on transient heat transfer measurement method. The changing of blowing ratio from 1.7 to 3.3 leads to film cooling effectiveness increasing on both pressure side and suction side. The Mach number or Reynolds number has no effect on the film cooling effectiveness on pressure side nearly, while increasing these two factors has opposite effect on film cooling performance on suction side. The increasing Mach number decreases the film cooling effectiveness at the rear region mainly, while at higher Reynolds number condition, the whole suction surface has significantly higher film cooling effectiveness because of the increasing cooling air mass flow rate. When changing the incidence angle from −15° to 0°, the film cooling effectiveness of pressure side decreases, and it presents the opposite trend on suction side. At off-design incidence of −15° and −10°, there is a low peak following the leading edge on the pressure side caused by the separation bubble, but it disappears with the incidence and blowing ratio increased.

Sign in / Sign up

Export Citation Format

Share Document