Heat Transfer Coefficients on the Squealer Tip and Near-Tip Regions of a Gas Turbine Blade With Single or Double Squealer

2003 ◽  
Vol 125 (4) ◽  
pp. 778-787 ◽  
Author(s):  
Jae Su Kwak ◽  
Jaeyong Ahn ◽  
Je-Chin Han ◽  
C. Pang Lee ◽  
Ronald S. Bunker ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blade Tip

Detailed heat transfer coefficient distributions on a gas turbine squealer tip blade were measured using a hue detection based transient liquid-crystals technique. The heat transfer coefficients on the shroud and near tip regions of the pressure and suction sides of a blade were also measured. Squealer rims were located along (a) the camber line, (b) the pressure side, (c) the suction side, (d) the pressure and suction sides, (e) the camber line and the pressure side, and (f) the camber line and the suction side, respectively. Tests were performed on a five-bladed linear cascade with a blow down facility. The Reynolds number based on the cascade exit velocity and the axial chord length of a blade was 1.1×106 and the overall pressure ratio was 1.2. Heat transfer measurements were taken at the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span. Results show that the heat transfer coefficients on the blade tip and the shroud were significantly reduced by using a squealer tip blade. Results also showed that a different squealer geometry arrangement changed the leakage flow path and resulted in different heat transfer coefficient distributions. The suction side squealer tip provided the lowest heat transfer coefficient on the blade tip and near tip regions compared to the other squealer geometry arrangements.

Heat Transfer Coefficients on the Squealer Tip and Near Tip Regions of a Gas Turbine Blade With Single or Double Squealer

2003 ◽  
Author(s):  
Jae Su Kwak ◽  
Jaeyong Ahn ◽  
Je-Chin Han ◽  
C. Pang Lee ◽  
Robert Boyle ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blade Tip

Detailed heat transfer coefficient distributions on a gas turbine squealer tip blade were measured using a hue detection based transient liquid crystals technique. The heat transfer coefficients on the shroud and near tip regions of the pressure and suction sides of a blade were also measured. Squealer rims were located along (a) the camber line, (b) the pressure side, (c) the suction side, (d) the pressure and suction sides, (e) the camber line and the pressure side, and (f) the camber line and the suction side, respectively. Tests were performed on a five-bladed linear cascade with a blow down facility. The Reynolds number based on the cascade exit velocity and the axial chord length of a blade was 1.1×106 and the overall pressure ratio was 1.2. Heat transfer measurements were taken at the three tip gap clearances of 1.0%, 1.5% and 2.5% of blade span. Results show that the heat transfer coefficients on the blade tip and the shroud were significantly reduced by using a squealer tip blade. Results also showed that a different squealer geometry arrangement changed the leakage flow path and resulted in different heat transfer coefficient distributions. The suction side squealer tip provided the lowest heat transfer coefficient on the blade tip and near tip regions compared to the other squealer geometry arrangements.

Effect of Squealer Geometry Arrangement on a Gas Turbine Blade Tip Heat Transfer

2002 ◽  
Vol 124 (3) ◽  
pp. 452-459 ◽  
Author(s):  
Gm Salam Azad ◽  
Je-Chin Han ◽  
Ronald S. Bunker ◽  
C. Pang Lee
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Gas Turbine Blade ◽  
Blade Tip

This study investigates the effect of a squealer tip geometry arrangement on heat transfer coefficient and static pressure distributions on a gas turbine blade tip in a five-bladed stationary linear cascade. A transient liquid crystal technique is used to obtain detailed heat transfer coefficient distribution. The test blade is a linear model of a tip section of the GE E3 high-pressure turbine first stage rotor blade. Six tip geometry cases are studied: (1) squealer on pressure side, (2) squealer on mid camber line, (3) squealer on suction side, (4) squealer on pressure and suction sides, (5) squealer on pressure side plus mid camber line, and (6) squealer on suction side plus mid camber line. The flow condition during the blowdown tests corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1×106. Results show that squealer geometry arrangement can change the leakage flow and results in different heat transfer coefficients to the blade tip. A squealer on suction side provides a better benefit compared to that on pressure side or mid camber line. A squealer on mid camber line performs better than that on a pressure side.

Effect of Squealer Geometry Arrangement on Gas Turbine Blade Tip Heat Transfer

2001 ◽  
Author(s):  
Gm Salam Azad ◽  
Je-Chin Han ◽  
Ronald S. Bunker ◽  
C. Pang Lee
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Gas Turbine Blade ◽  
Blade Tip

Abstract This study investigates the effect of a squealer tip geometry arrangement on heat transfer coefficient and static pressure distributions on a gas turbine blade tip in a five-bladed stationary linear cascade. A transient liquid crystal technique is used to obtain detailed heat transfer coefficient distribution. The test blade is a linear model of a tip section of the GE E3 high-pressure turbine first stage rotor blade. Six tip geometry cases are studied: 1) squealer on pressure side, 2) squealer on mid camber line, 3) squealer on suction side, 4) squealer on pressure and suction sides, 5) squealer on pressure side plus mid camber line, and 6) squealer on suction side plus mid camber line. The flow condition corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1 × 106. Results show that squealer geometry arrangement can change the leakage flow and results in different heat transfer coefficients to the blade tip. A squealer on suction side provides a better benefit compared to that on pressure side or mid camber line. A squealer on mid camber line performs better than that on a pressure side.

Effect of Tip and Pressure Side Coolant Injection on Heat Transfer Distributions for a Plane and Recessed Tip

2005 ◽  
Vol 129 (1) ◽  
pp. 151-163 ◽  
Author(s):  
Hasan Nasir ◽  
Srinath V. Ekkad ◽  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Pressure Side ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio ◽  
Coolant Injection ◽  
Blade Tip ◽  
Lift Off

The present study investigates the effects of coolant injection on adiabatic film effectiveness and heat transfer coefficients from a plane and recessed tip of a high pressure turbine first stage rotor blade. Three cases where coolant is injected from (a) five orthogonal holes located along the camber line, (b) seven angled holes located near the blade tip along the pressure side, and (c) combination cases when coolant is injected from both tip and pressure side holes were studied. The pressure ratio (inlet total pressure to exit static pressure for the cascade) across the blade row was 1.2, and the experiments were run in a blow-down test rig with a four-blade linear cascade. The Reynolds number based on cascade exit velocity and axial chord length was 8.61×105 and the inlet and exit Mach numbers were 0.16 and 0.55, respectively. A transient infrared technique was used to measure adiabatic film effectiveness and heat transfer coefficient simultaneously for three blowing ratios of 1.0, 2.0, and 3.0. For all the cases, gap-to-blade span ratio of 1% was used. The depth-to-blade span ratio of 0.0416 was used for the recessed tip. Pressure measurements on the shroud were also taken to characterize the leakage flow and understand the heat transfer distributions. For tip injection, when blowing ratio increases from 1.0 to 2.0, film effectiveness increases for both plane and recessed tip and heat transfer coefficient decreases for both plane and recessed tip. At blowing ratio 3.0, lift-off is observed for both cases. In case of pressure side coolant injection and for plane tip, lift-off is observed at blowing ratio 2.0 and reattachments of jets are observed at blowing ratio 3.0. But, almost no effectiveness is observed for squealer tip at all blowing ratios with pressure side injection with reduced heat transfer coefficient along the pressure side. For combination case, very high effectiveness is observed at blowing ratio 3.0 for both plane and recessed blade tip. It appears that for this high blowing ratio, coolant jets from the tip hit the shroud first and then reattach back onto the blade tip with very high heat transfer coefficients for both plane and recessed blade tip.

Heat Transfer Coefficient and Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2002 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Rotor Blade ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

The detailed distributions of heat transfer coefficient and film cooling effectiveness on a gas turbine blade tip were measured using a hue detection based transient liquid crystal technique. Tests were performed on a five-bladed linear cascade with blow down facility. The blade was a 2-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The Reynolds number based on cascade exit velocity and axial chord length was 1.1 × 106 and the total turning angle of the blade was 97.7°. The overall pressure ratio was 1.32 and the inlet and exit Mach number were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. The blade model was equipped with a single row of film cooling holes at both the tip portion along the camber line and near the tip region of the pressure-side. All measurements were made at the three different tip gap clearances of 1%, 1.5%, and 2.5% of blade span and the three blowing ratios of 0.5, 1.0, and 2.0. Results showed that, in general, heat transfer coefficient and film effectiveness increased with increasing tip gap clearance. As blowing ratio increased, heat transfer coefficient decreased, while film effectiveness increased. Results also showed that adding pressure-side coolant injection would further decrease blade tip heat transfer coefficient but increase film effectiveness.

scholarly journals Endwall Heat Transfer in a Vane Cascade Passage and in a Curved Duct

1989 ◽  
Author(s):  
M. T. Boyle ◽  
K. V. Hoose
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Suction Side ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Steel Strips ◽  
Surface Heat Transfer Coefficient ◽  
Duct Geometry

Measurements of endwall heat transfer coefficient have been made for flow through a vane cascade passage and a similarly shaped duct. The purpose of this work is to evaluate the usefulness of the duct shape for modelling cascade endwall heat transfer. The Reynolds number for the cascade experiment is 5.4 × 105 based on chord. For the duct geometry the inlet velocity is adjusted in order to match the Reynolds number based on the pitch dimension. Thin stainless steel strips mounted on the endwall make up a large flat resistance heater. The temperature distribution over the heater surface is measured with 156 resistance temperature sensors. Surface heat transfer coefficient is obtained in the vicinity of each sensor by a local energy balance. The results are presented so as to show clearly the effect of secondary flow on endwall heat transfer. Except in the entrance plane region, the qualities of the cascade endwall Stanton number distribution compare well with the duct endwall measurements. The duct endwall heat transfer coefficients are everywhere greater than for the cascade passage endwall. The effect of a counter-clockwise rotating vortex in the suction side corner is shown clearly for both geometries.

Heat Transfer Coefficient on the Squealer Tip and Near Squealer Tip Regions of a Gas Turbine Blade

2002 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Rotor Blade ◽  
Intensity Level ◽  
Gas Turbine Blade ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients

Detailed heat transfer coefficient distributions on a squealer tip of a gas turbine blade were measured using a hue detection based transient liquid crystals technique. The heat transfer coefficients on the shroud and near tip region of the pressure and suction sides of a blade were also measured. Tests were performed on a five-bladed linear cascade with blow down facility. The blade was a 2-dimensional model of a first stage gas turbine rotor blade with a profile of a GE-E3 aircraft gas turbine engine rotor blade. The Reynolds number based on the cascade exit velocity and axial chord length of a blade was 1.1×106 and the total turning angle of the blade was 97.7°. The overall pressure ratio was 1.23 and the inlet and exit Mach number were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. The heat transfer measurements were taken at the three different tip gap clearances of 1.0%, 1.5% and 2.5% of blade span. Results showed that the overall heat transfer coefficient on the squealer tip was higher than that on the shroud and the near tip region of the pressure and suction side. Results also showed that the heat transfer coefficients on the squealer tip and its shroud were lower than that on the plane tip and shroud, but the heat transfer coefficients on the near tip region of suction and pressure sides were higher for the squealer tip case.

Heat Transfer and Flow on the First-Stage Blade Tip of a Power Generation Gas Turbine: Part 1—Experimental Results

1999 ◽  
Vol 122 (2) ◽  
pp. 263-271 ◽  
Author(s):  
Ronald S. Bunker ◽  
Jeremy C. Bailey ◽  
Ali A. Ameri
Keyword(s):  
Heat Transfer ◽  
Power Generation ◽  
Heat Transfer Coefficient ◽  
Mach Number ◽  
Transfer Coefficient ◽  
Pressure Distribution ◽  
Turbulence Intensity ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blade Tip

A combined experimental and computational study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first-stage blade tip surface for a geometry typical of large power generation turbines (>100 MW). This paper is concerned with the design and execution of the experimental portion of the study, which represents the first reported investigation to obtain nearly full surface information on heat transfer coefficients within an environment that develops an appropriate pressure distribution about an airfoil blade tip and shroud model. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high-pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57×106, and total turning of about 110 deg. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5 or 9 percent. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp edge and rounded edge tip geometries at each of the inlet turbulence intensity levels. [S0889-504X(00)01902-4]

Heat Transfer Augmentation Downstream of Rows of Various Dimple Geometries on the Suction Side of a Gas Turbine Airfoil

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Jason E. Dees ◽  
David G. Bogard ◽  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Transfer Coefficients ◽  
Film Cooling Holes

Heat transfer coefficients were measured downstream of a row of shaped film cooling holes, as well as elliptical, diffuser, and teardrop shaped dimples, simulating depressions due to film coolant holes of different shapes. These features were placed on the suction side of a simulated gas turbine vane. The dimples were used as approximations to film cooling holes after the heat transfer levels downstream of active fan shaped film cooling holes was found to be independent of film cooling. The effects of the dimples were tested with varying approach boundary layers, freestream turbulence intensity, and Reynolds numbers. For the case of an untripped (transitional) approach boundary layer, all dimple shapes caused approximately a factor of 2 increase in heat transfer coefficient relative to the smooth baseline condition due to the dimples effectively causing boundary layer transition downstream. The exact augmentation varied depending on the dimple geometry: diffuser shapes causing the largest augmentation and teardrop shapes causing the lowest augmentation. For tripped (turbulent boundary layer) approach conditions, the dimple shapes all caused the same 20% augmentation relative to the smooth tripped baseline. The already turbulent nature of the tripped approach flow reduces the effect that the dimples have on the downstream heat transfer coefficient.

Heat Transfer Augmentation Downstream of Rows of Various Dimple Geometries on the Suction Side of a Gas Turbine Airfoil

2008 ◽  
Author(s):  
Jason E. Dees ◽  
David G. Bogard ◽  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Transfer Coefficients ◽  
Film Cooling Holes

Heat transfer coefficients were measured downstream of a row of shaped film cooling holes as well as elliptical, diffuser, and teardrop shaped dimples simulating depressions due to film coolant holes of different shapes. These features were placed on the suction side of a simulated gas turbine vane. The dimples were used as approximations to film cooling holes after the heat transfer levels downstream of active fan shaped film cooling holes was found to be independent of film cooling. The effects of the dimples were tested with varying approach boundary layers, freestream turbulence intensity, and Reynolds numbers. For the case of an untripped (transitional) approach boundary layer, all dimple shapes caused approximately a factor of two increase in heat transfer coefficient relative to the smooth baseline condition due to the dimples effectively causing boundary layer transition downstream. The exact augmentation varied depending on the dimple geometry: diffuser shapes causing the largest augmentation and teardrop shapes causing the lowest augmentation. For tripped (turbulent boundary layer) approach conditions, the dimple shapes all caused the same 20% augmentation relative to the smooth tripped baseline. The already turbulent nature of the tripped approach flow reduces the effect that the dimples have on the downstream heat transfer coefficient.

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