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

Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film-cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. 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 test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film-cooling holes on the pressure-side near the tip region and the tip surface along the camber line. A hue detection based transient liquid crystal technique was used to measure heat transfer coefficients and film-cooling effectiveness. All measurements were done for the tip gap clearances of 1.0%,1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1 × 106 and the overall pressure ratio was 1.32. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film-cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film-cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

Journal of Turbomachinery ◽

10.1115/1.1622712 ◽

2003 ◽

Vol 125 (4) ◽

pp. 648-657 ◽

Cited By ~ 31

Author(s):

Jae Su Kwak ◽

Je-Chin Han

Keyword(s):

Heat Transfer ◽

Gas Turbine ◽

Turbine Blade ◽

Film Cooling ◽

Rotor Blade ◽

Transfer Coefficients ◽

Cooling Effectiveness ◽

Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a two-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film cooling holes on the pressure side near the tip region and the tip surface along the camber line. Hue detection based transient liquid crystals technique was used to measure heat transfer coefficients and film cooling effectiveness. All measurements were done for the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. 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 97.9 deg. The overall pressure ratio was 1.2 and the inlet and exit Mach numbers were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

Heat Transfer and Film Cooling on a Contoured Blade Endwall With Platform Gap Leakage

Journal of Turbomachinery ◽

10.1115/1.4035202 ◽

2017 ◽

Vol 139 (5) ◽

Cited By ~ 4

Author(s):

Stephen P. Lynch ◽

Karen A. Thole

Keyword(s):

Heat Transfer ◽

Turbine Blade ◽

Film Cooling ◽

Vortical Flow ◽

Leakage Flow ◽

Transfer Coefficients ◽

Cooling Effectiveness ◽

Manufacturing Assembly

Turbine blade components in an engine are typically designed with gaps between parts due to manufacturing, assembly, and operational considerations. Coolant is provided to these gaps to limit the ingestion of hot combustion gases. The interaction of the gaps, their leakage flows, and the complex vortical flow at the endwall of a turbine blade can significantly impact endwall heat transfer coefficients and the effectiveness of the leakage flow in providing localized cooling. In particular, a platform gap through the passage, representing the mating interface between adjacent blades in a wheel, has been shown to have a significant effect. Other important turbine blade features present in the engine environment are nonaxisymmetric contouring of the endwall, and an upstream rim seal with a gaspath cavity, which can reduce and increase endwall vortical flow, respectively. To understand the platform gap leakage effect in this environment, measurements of endwall heat transfer, and film cooling effectiveness were performed in a scaled blade cascade with a nonaxisymmetric contour in the passage. A rim seal with a cavity, representing the overlap interface between a stator and rotor, was included upstream of the blades and a nominal purge flowrate of 0.75% of the mainstream was supplied to the rim seal. The results indicated that the endwall heat transfer coefficients increased as the platform gap net leakage increased from 0% to 0.6% of the mainstream flowrate, but net heat flux to the endwall was reduced due to high cooling effectiveness of the leakage flow.

Numerical Simulation of Film Cooling on the Tip of a Gas Turbine Blade

Volume 3: Turbo Expo 2002, Parts A and B ◽

10.1115/gt2002-30553 ◽

2002 ◽

Cited By ~ 15

Author(s):

Sumanta Acharya ◽

Huitao Yang ◽

Srinath V. Ekkad ◽

Chander Prakash ◽

Ron Bunker

Keyword(s):

Heat Transfer ◽

Turbine Blade ◽

Film Cooling ◽

Pressure Ratio ◽

Transfer Coefficients ◽

Local Pressure ◽

Blowing Ratio ◽

Coolant Delivery

Numerical simulations of flow and heat transfer are presented for a GE-E3 turbine blade with a film-cooled tip. Results are presented for both a flat tip and a squealer tip. Straight-through coolant holes are considered, and the calculation domain includes the flow development in the coolant delivery tubes. Results are presented with three different tip gaps representing 1%, 1.5% and 2.5% of blade span, a blowing ratio (ratio of coolant-jet-exit velocity to average passage flow velocity) of 1, and an inlet turbulence intensity of 6.1%. On a flat tip, film coolant injection is shown to lower the local pressure ratio and alters the nature of the leakage vortex. High film cooling effectiveness and low heat transfer coefficients are obtained along the coolant trajectory; these values increase slightly with increasing tip clearances. For a squealer tip, the flow inside the squealer cavity exhibits streamwise directed flow, which alters the trajectory of the coolant jets and reduces their effectiveness.

Film-Cooling Characteristics of Pressure-Side Discharge Slots in an Accelerating Mainstream Flow

Volume 3: Turbo Expo 2005, Parts A and B ◽

10.1115/gt2005-69061 ◽

2005 ◽

Cited By ~ 7

Author(s):

Yong W. Kim ◽

Chad Coon ◽

Hee-Koo Moon

Keyword(s):

Heat Transfer ◽

Film Cooling ◽

Local Heat Transfer ◽

Local Heat ◽

Transfer Coefficients ◽

Cooling Effectiveness ◽

Blowing Ratio ◽

Mainstream Flow

Pressure-side discharge is commonly employed in turbine blades and nozzle guide vanes to keep the trailing edge metal temperatures within an allowable limit while minimizing aerodynamic penalties. Despite its widespread use, film-cooling data of the discharge slot are scarce in open literature. The objectives of the present experimental study were to measure detailed local heat transfer and film-cooling effectiveness from a 10x scale trailing-edge model of an industrial gas turbine airfoil in a low speed wind tunnel. To simulate the mainstream flow acceleration in vane and blade row passages, a linear velocity gradient was imposed using an adjustable top wall. The present work employed the composite slab quasi-steady liquid crystal method that allows measurements of local heat transfer coefficients and film-cooling effectiveness from two related tests. With this technique, the heat transfer measurement can be performed in a cold wind tunnel. The coolant-to-mainstream blowing ratio was varied between 0.25 and 1.0. The slot hydraulic diameter based Reynolds number ranged from 4,760 to 19,550. The coolant-to-mainstream density ratio was fixed at 0.95. Slot discharge coefficients were also measured with mainstream acceleration. Both local heat transfer coefficients and film-cooling effectiveness displayed a strong dependency on blowing ratio and mainstream acceleration. However, the discharge coefficients showed little dependency on the mainstream acceleration.

Heat Transfer and Film Cooling on a Contoured Blade Endwall With Platform Gap Leakage

Volume 5B: Heat Transfer ◽

10.1115/gt2015-43301 ◽

2015 ◽

Cited By ~ 3

Author(s):

Stephen P. Lynch ◽

Karen A. Thole

Keyword(s):

Heat Transfer ◽

Turbine Blade ◽

Film Cooling ◽

Vortical Flow ◽

Leakage Flow ◽

Transfer Coefficients ◽

Cooling Effectiveness ◽

Manufacturing Assembly

Turbine blade components in an engine are typically designed with gaps between parts due to manufacturing, assembly, and operational considerations. Coolant is provided to these gaps to limit the ingestion of hot combustion gases. The interaction of the gaps, their leakage flows, and the complex vortical flow at the endwall of a turbine blade can significantly impact endwall heat transfer coefficients and the effectiveness of the leakage flow in providing localized cooling. In particular, a platform gap through the passage, representing the mating interface between adjacent blades in a wheel, has been shown to have a significant effect. Other important turbine blade features present in the engine environment are non-axisymmetric contouring of the endwall, and an upstream rim seal with a gaspath cavity, which can reduce and increase endwall vortical flow, respectively. To understand the platform gap leakage effect in this environment, measurements of endwall heat transfer and film cooling effectiveness were performed in a scaled blade cascade with a non-axisymmetric contour in the passage. A rim seal with a cavity, representing the overlap interface between a stator and rotor, was included upstream of the blades and a nominal purge flowrate of 0.75% of the mainstream was supplied to the rim seal. Results indicated that endwall heat transfer coefficients increased as platform gap net leakage increased from 0% to 0.6% of the mainstream flowrate, but net heat flux to the endwall was reduced due to high cooling effectiveness of the leakage flow.

Detailed Film Cooling Measurements over a Gas Turbine Blade Using a Transient Liquid Crystal Image Technique

International Journal of Rotating Machinery ◽

10.1155/s1023621x01000367 ◽

2001 ◽

Vol 7 (6) ◽

pp. 415-424 ◽

Cited By ~ 1

Author(s):

Hui Du ◽

Srinath V. Ekkad ◽

Je-Chin Han ◽

C. Pang Lee

Keyword(s):

Heat Transfer ◽

Liquid Crystal ◽

Gas Turbine ◽

Turbine Blade ◽

Film Cooling ◽

Gas Turbine Blade ◽

Transfer Coefficients ◽

Blowing Ratio ◽

Image Technique

Detailed heat transfer coefficient and film effectiveness distributions over a gas turbine blade with film cooling are obtained using a transient liquid crystal image technique. The test blade has three rows of film holes on the leading edge and two rows each on the pressure and suction surfaces. A transient liquid crystal technique maps the entire blade midspan region, and helps provide detailed measurements, particularly near the film hole. Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity is5.3×105. Two different coolants (air andCo2) were used to simulate coolant density effect. Coolant blowing ratio was varied between 0.8 and 1.2 for air injection and 0.4–1.2 forCo2injection. Results show that film injection promotes earlier laminar-turbulent boundary layer transition on the suction surface and also enhances local heat transfer coefficients (up to 80%) downstream of injection. An increase in coolant blowing ratio produces higher heat transfer coefficients for both coolants. This effect is stronger immediately downstream of injection holes. Film effectiveness is highest at a blowing ratio of 0.8 for air injection and at a blowing ratio of 1.2 forCo2injection. Such detailed results will help provide insight into the film cooling phenomena on a gas turbine blade.

Experimental and Numerical Investigation of Effect of Blowing Ratio on Film Cooling Effectiveness and Heat Transfer Coefficient Over a Gas Turbine Blade Leading Edge Film Cooling Configurations

ASME 2013 Gas Turbine India Conference ◽

10.1115/gtindia2013-3552 ◽

2013 ◽

Author(s):

Yepuri Giridhara Babu ◽

Gururaj Lalgi ◽

Ashok Babu Talanki Puttarangasetty ◽

Jesuraj Felix ◽

Sreenivas Rao V. Kenkere ◽

...

Keyword(s):

Heat Transfer ◽

Heat Transfer Coefficient ◽

Transfer Coefficient ◽

Gas Turbine ◽

Turbine Blade ◽

Film Cooling ◽

Gas Turbine Blade ◽

Cooling Effectiveness ◽

Blowing Ratio

Film cooling is one of the cooling techniques to cool the hot section components of a gas turbine engines. The gas turbine blade leading edges are the vital parts in the turbines as they are directly hit by the hot gases, hence the optimized cooling of gas turbine blade surfaces is essential. This study aims at investigating the film cooling effectiveness and heat transfer coefficient experimentally and numerically for the three different gas turbine blade leading edge models each having the one row of film cooling holes at 15, 30 and 45 degrees hole orientation angle respectively from stagnation line. Each row has the five holes with the hole diameter of 3mm, pitch of 20mm and has the hole inclination angle of 20deg. in spanwise direction. Experiments are carried out using the subsonic cascade tunnel facility of National Aerospace Laboratories, Bangalore at a nominal flow Reynolds number of 1,00,000 based on the leading edge diameter, varying the blowing ratios of 1.2, 1.50, 1.75 and 2.0. In addition, an attempt has been made for the film cooling effectiveness using CFD simulation, using k-€ realizable turbulence model to solve the flow field. Among the considered 15, 30 and 45 deg. models, both the cooling effectiveness and heat transfer coefficient shown the increase with the increase in hole orientation angle from stagnation line. The film cooling effectiveness increases with the increase in blowing ratio upto 1.5 for the 15 and 30 deg. models, whereas on the 45 deg. model the increase in effectiveness shown upto the blowing ratio of 1.75. The heat transfer coefficient values showed the increase with the increase in blowing ratio for all the considered three models. The CFD results in the form of temperature, velocity contours and film cooling effectiveness values have shown the meaningful results with the experimental values.

Simulation Film Cooling in the Leading Edge Region of a Turbine Blade (Trench Effect on Film Effectiveness from Cylinder in Crossflow)

Applied Mechanics and Materials ◽

10.4028/www.scientific.net/amm.554.317 ◽

2014 ◽

Vol 554 ◽

pp. 317-321

Author(s):

Mohamad Rasidi Bin Pairan ◽

Norzelawati Binti Asmuin ◽

Hamidon bin Salleh

Keyword(s):

Gas Turbine ◽

Turbine Blade ◽

Film Cooling ◽

Leading Edge ◽

Edge Region ◽

Cooling Effectiveness ◽

Blowing Ratio ◽

Hot Gas ◽

Cooling Technique

Film cooling is one of the cooling techniques applied to the turbine blade. Gas turbine used film cooling technique to protect turbine blade from directly expose to the hot gas to avoid the blade from defect. The focus of this investigation is to investigate the effect of embedded three difference depth of trench at coolant holes geometry. Comparisons are made at four difference blowing ratios which are 1.0, 1.25 and 1.5. Three configuration leading edge with depth Case A (0.0125D), Case B (0.0350D) and Case C (0.713D) were compared to leading edge without trench. Result shows that as blowing ratio increased from 1.0 to 1.25, the film cooling effectiveness is increase for leading edge without trench and also for all cases. However when the blowing ratio is increase to 1.5, film cooling effectiveness is decrease for all cases. Overall the Case B with blowing ratio 1.25 has the best film cooling effectiveness with significant improvement compared to leading edge without trench and with trench Case A and Case C.

Effect of Surface Roughness on Local Heat Transfer and Film Cooling Effectiveness

Volume 4: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/95-gt-014 ◽

1995 ◽

Cited By ~ 17

Author(s):

Douglas N. Barlow ◽

Yong W. Kim

Keyword(s):

Heat Transfer ◽

Film Cooling ◽

Smooth Surface ◽

Rough Surfaces ◽

Local Heat Transfer ◽

Local Heat ◽

Transfer Coefficients ◽

Cooling Effectiveness ◽

Heat Transfer Coefficients

An experimental investigation of film cooling on rough surfaces has been accomplished at a Reynolds number and dimensionless boundary layer momentum thickness found in current high performance first stage turbine vanes. A transient experimental method using thermochromic liquid crystals is employed to determine both local heat transfer coefficients and film cooling effectiveness values on planar rough surfaces. Two surface roughness configurations are investigated with a single row of cooling holes spaced three diameters apart and inclined 30° to the mainstream flow. The mainstream turbulence level at the point of film injection is 8.5% and the density ratio considered is approximately 1.0. The influence of roughness on the centerline film cooling effectiveness, laterally averaged film cooling effectiveness, laterally averaged heat transfer coefficients, as well as area averaged values are presented. It is found that the presence of roughness causes a decrease in the film cooling effectiveness over that of the smooth surface for the range of experimental parameters considered in this study. In addition, significant lateral smoothing in film cooling effectiveness distribution is observed for the rougher surfaces. Measured heat transfer coefficients on rough surfaces show a trend of monotonic increase with blowing ratio. However, such increase is not as great as that for the case of smooth surface.