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

2005 ◽  
Author(s):  
Hasan Nasir ◽  
Srinath V. Ekkad ◽  
Ronald S. Bunker
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Coolant Injection ◽  
Blade Tip ◽  
Lift Off ◽  
Tip Injection ◽  
Very High

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 HPT 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 number were 0.16 and 0.55, respectively. A transient infrared (IR) 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. 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. 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 on to the blade tip. For tip injection, as blowing ratio increases heat transfer coefficient decreases for both plane and recessed tip. In case of pressure side coolant injection and for plane tip, film injection reduced heat transfer coefficient along the pressure side. Minimal effect is observed for recessed tip at all blowing ratios. For combination case, very high heat transfer coefficient 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 on to the blade tip.

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.

Blade Tip Leakage Flow and Heat Transfer With Pressure-Side Winglet

2003 ◽  
Author(s):  
Arun K. Saha ◽  
Sumanta Acharya ◽  
Chander Prakash ◽  
Ron Bunker
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Leakage Flow ◽  
Pressure Side ◽  
Average Heat Transfer Coefficient ◽  
Tip Leakage Flow ◽  
Average Heat ◽  
Flow And Heat Transfer ◽  
Blade Tip

A numerical study has been conducted to explore the effect of a pressure-side winglet on the flow and heat transfer over a blade tip. Calculations are performed for both a flat tip and a squealer tip. The winglet is in the form of a flat extension, and is shaped in the axial chord direction to have the maximum thickness at the chord location where the pressure difference is the largest between the pressure and suction sides. For the flat tip, the pressure side winglet exhibits a significant reduction in the leakage flow strength and an associated reduction in the aerodynamic loss. The low heat transfer coefficient “sweet-spot” region is larger with the pressure-side winglet, and lower heat transfer coefficients are also observed along the pressure side of the blade. The winglet reduces the average heat transfer coefficient by about 7%. In the presence of a squealer, the role of the winglet decreases significantly, and only a 0.5% reduction in the pressure ratio is achieved with the winglet with virtually no reduction in the average heat transfer coefficient.

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.

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 casing purge flow on heat transfer and cooling performance of blade squealer tip for a gas turbine stage

2021 ◽  
pp. 095765092110403
Author(s):  
Shijie Jiang ◽  
Zhigang Li ◽  
Jun Li
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Three Dimensional ◽  
Tip Clearance ◽  
Local Cooling ◽  
Cold Air ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Blade Tip

The first stage of GE-E3 turbine is employed to investigate effect of casing purge flow upstream rotor blade tip. Three-dimensional Reynolds-averaged Navier-Stokes (RANS) equations and standard k-ω model are solved to obtain tip heat transfer simulations. The results reveal that: heat transfer coefficient of blade tip surface can be significantly reduced when casing purge flow is set. Tip averaged heat transfer coefficient of cases with and without swirly velocity casing purge flow decrease 3.5% and 3.4% compared with the case without casing purge flow. Compared with case which blowing ratio equals to 0.5, it can be found that averaged tip heat transfer coefficient of cases which blowing ratio equals to 1.0 and 1.5 decrease 2.3% and 1.8%, respectively. Setting blowing ratio as 1.0 can best cool tip surface without wasting cold air resources. Increasing rotating speed can induce cold air entering tip trailing region and improve local cooling effect. Flow structure inside the tip clearance are also revealed and discussed.

Effects of Oscillations in the Mainstream on Film Cooling at Various Blowing Ratios

2017 ◽  
Author(s):  
Seung Il Baek ◽  
Savas Yavuzkurt
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Jet Flow ◽  
Main Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Lift Off

The objective of this study is to understand the effects of flow oscillations in the mainstream and film cooling jets on film cooling at various blowing ratios (0.5, 0.78, 1.0 and 1.5). These oscillations could be caused by the combustion instabilities. They are approximated in sinusoidal form for the current study. The effects of different frequencies (0, 2, 16, 32 Hz) on film cooling are investigated. Simulations are performed using URANS Realizable k-epsilon and LES Smagorinsky-Lilly turbulence models. The results indicate that if the frequencies of the mainstream and the jet flow are increased at a low average blowing ratio of M = 0.5, the adiabatic film cooling effectiveness is decreased and the heat transfer coefficient is increased due to increased disturbance in jet and main flow interaction with increasing frequency. It was observed that when the frequency of the mainstream and the cooling jet flow is increased at M = 0.5, the amplitude of the pressure difference between the mainstream and the plenum is increased resulting in increased amplitude of coolant flow rate oscillations leading to more jet lift off and more disturbance in the main flow and coolant interaction. Consequently, adiabatic film cooling effectiveness is decreased and heat transfer coefficient is increased. If the frequency of the mainstream is increased from 0 Hz to 2, 16, or 32 Hz at M = 0.5, the centerline effectiveness is decreased about 10%, 12%, or 47% and the spanwise-averaged Stanton number ratio is increased about 4%, 5%, or 9% respectively. If the frequencies of the main flow and the jet flow are increased at higher blowing ratios of M = 1.0 and 1.5, adiabatic effectiveness is increased and the spanwise-averaged heat transfer coefficient are decreased. Under steady flow conditions jet lift off is generated for these high blowing ratios. If the frequency of the mainstream and the jet flow is increased, the amplitude of coolant jet flow rate oscillation is increased for the same reason as mentioned above for M = 0.5. This leads to less jet lift off during the cycle resulting in more frequent coolant contact with the wall and consequently increased centerline effectiveness as frequency increases. In addition, the entrainment of hot gases underneath the jet doesn’t lead to higher mixing between the hot mainstream and the coolant and this results in decreased heat transfer coefficient. This is also indicated by the turbulent kinetic energy levels. Some representative results are: when the frequency of the main flow is increased from 0 Hz to 2, 16, or 32 Hz at M = 1.0, the centerline effectiveness is increased about 8%, 19%, or 320%. Also, if the oscillation frequency is increased from 2 Hz to 16, or 32 Hz at M = 1.0, the spanwise-averaged Stanton number ratio is decreased around 2%, to 5% respectively. It seems like the cut off point for low and high blowing ratio behavior of cooling jets is around M = 0.78.

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.

Heat transfer characteristics analysis on a fully cooled vane with varied density ratios

2021 ◽  
pp. 1-19
Author(s):  
Chunyi Yao ◽  
Zheng Zhang ◽  
Bo-lun Zhang ◽  
Hui Ren Zhu ◽  
Cun Liang Liu
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Coolant Injection ◽  
Density Ratios

Abstract The objective of this experimental investigation was to determine the cooling performance of a fully cooled vane with 18 rows of cylinder cooling holes. The exit Reynolds number in the wind tunnel normalized by the true chord was 500,000 with a turbulence intensity of 15%. The film cooling effectiveness and heat transfer coefficient distribution were obtained by the transient liquid crystal technology, three mass flow ratios (MFR=7.0%, 9.9%, 11%) and two density ratios (DR=1.0, 1.5) were tested. The results show that the film cooling effectiveness distribution on the suction side is more uniform and the coolant injection trajectory is much longer than that on the pressure side. As the density ratio increasing to 1.5, the more laterally uniform film cooling effectiveness contour on the pressure side is observed and the spatially averaged film cooling effectiveness is increased by 11%∼43%. For the MFR=7.0%, the coolant injection with low momentum thickens the boundary layer and reduces the heat transfer coefficient in the mid-chord region of the pressure side. Both the increased mass flow ratio and decreased density ratio result in a higher heat transfer coefficient, while do not alter the distribution trend. By calculating the heat flux ratio, the reduction in the heat flux at DR=1.5 is found to be within 20% in most areas than that of DR=1.0 on the vane surface.

An Experimental Investigation of Adiabatic Film Cooling Effectiveness and Heat Transfer Coefficient on a Transonic Squealer Tip

2018 ◽  
Author(s):  
Andrew J. Saul ◽  
Peter T. Ireland ◽  
John D. Coull ◽  
Tsun Holt Wong ◽  
Haidong Li ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Mass Flow ◽  
Film Cooling ◽  
Leakage Flow ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

The effect of film cooling on a high pressure turbine blade with an open squealer tip has been examined in a high speed linear cascade. The cascade operates at engine realistic Mach and Reynolds numbers, producing transonic flow conditions over the blade tip. Tests have been performed on two uncooled tip geometries with differing pressure side rim edge radii, and a cooled tip matching one of the uncooled cases. The pressure sensitive paint technique has been used to measure adiabatic film cooling effectiveness on the blade tip at a range of tip gaps and coolant mass flow rates. Complementary tip heat transfer coefficients (HTC) have been measured using transient infrared thermography, and the effects of the coolant film on the tip heat transfer and engine heat flux examined. The uncooled data show that the tip heat transfer coefficient distribution is governed by the nature of flow reattachments and impingements. The squealer tip can be broken down into three regions, each exhibiting a distinct response to a change in the tip gap, depending on the local behaviour of the overtip leakage flow. The edge radius of the pressure side rim causes the overtip leakage flow to change dramatically at low clearance. Complementary CFD shows that the addition of casing motion causes no further change on the pressure side rim. Injected coolant interacts with the overtip leakage flow, which can locally enhance the tip heat transfer coefficient compared to the uncooled tip. The film effectiveness is dependent on both the coolant mass flow rate and tip clearance. At increased coolant mass flow, areas of high film effectiveness on the pressure side rim coincide strongly with a net heat flux reduction and in the subsonic tip region with low heat transfer coefficient.

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