The Effect of Combustor-Turbine Interface Gap Leakage on the Endwall Heat Transfer for a Nozzle Guide Vane

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
S. P. Lynch ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Guide Vane ◽  
Flux Ratio ◽  
Slot Width ◽  
Leakage Flow ◽  
Transfer Coefficients ◽  
Coverage Area ◽  
Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Nozzle Guide

To enable turbine components to withstand high combustion temperatures, they are cooled by air routed from the compressor, which can leak through gaps between components. These gaps vary in size from thermal expansions that take place. The leakage flow between the interface of the combustor and the turbine, in particular, interacts with the flowfield along the endwall. This study presents measurements of adiabatic cooling effectiveness and heat transfer coefficients on the endwall of a first vane, with the presence of leakage flow through a flush slot upstream of the vane. The effect of axial contraction of the slot width due to thermal expansion of the engine was tested for two blowing rates. Contracting the slot width, while maintaining the slot mass flow, resulted in a larger coolant coverage area and higher effectiveness values, as well as slightly lower heat transfer coefficients. Matching the momentum flux ratio of the leakage flow from the nominal and contracted slot widths lowered both cooling effectiveness and heat transfer coefficients for the contracted slot flow. Comparison of the coolant coverage pattern to the measured endwall shear stress topology indicated that the trajectory of the slot coolant was dictated by the complex endwall flow.

The Effect of Combustor-Turbine Interface Gap Leakage on the Endwall Heat Transfer for a Nozzle Guide Vane

2007 ◽  
Author(s):  
S. P. Lynch ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Guide Vane ◽  
Flux Ratio ◽  
Slot Width ◽  
Leakage Flow ◽  
Transfer Coefficients ◽  
Coverage Area ◽  
Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Nozzle Guide

To enable turbine components to withstand high combustion temperatures, they are cooled by air routed from the compressor, which can leak through gaps between components. These gaps vary in size from thermal expansions that take place. The leakage flow between the interface of the combustor and the turbine, in particular, interacts with the flowfield along the endwall. This study presents measurements of adiabatic cooling effectiveness and heat transfer coefficients on the endwall of a first vane, with the presence of leakage flow through a flush slot upstream of the vane. The effect of axial contraction of the slot width due to thermal expansion of the engine was tested for two blowing rates. Contracting the slot width, while maintaining the slot mass flow, resulted in a larger coolant coverage area and higher effectiveness values, as well as slightly lower heat transfer coefficients. Matching the momentum flux ratio of the leakage flow from the nominal and contracted slot widths lowered both cooling effectiveness and heat transfer coefficients for the contracted slot flow. Comparison of the coolant coverage pattern to the measured endwall shear stress topology indicated that the trajectory of the slot coolant was dictated by the complex endwall flow.

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

2017 ◽  
Vol 139 (5) ◽  
Author(s):  
Stephen P. Lynch ◽  
Karen A. Thole
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Vortical Flow ◽  
Leakage Flow ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
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.

The Influence of the Boundary Layer State and Reynolds Number on Film Cooling and Heat Transfer on a Cooled Nozzle Guide Vane

2000 ◽  
Author(s):  
Hans Reiss ◽  
Albin Bölcs
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Flow Conditions ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Film cooling and heat transfer measurements were carried out on a cooled nozzle guide vane in a linear cascade, using a transient liquid crystal technique. Three flow conditions were realized: the nominal operating condition of the vane with an exit Reynolds number of 1.47e6, as well as two lower flow conditions: Re2L = 1.0e6 and 7.5e5. The vane model was equipped with a single row of inclined round film cooling holes with compound angle orientation on the suction side. Blowing ratios ranging form 0.3 to 1.5 were covered, all using foreign gas injection (CO2) yielding an engine-representative density ratio of 1.6. Two distinct states of the incoming boundary layer onto the injection station were compared, an undisturbed laminar boundary layer as it forms naturally on the suction side, and a fully turbulent boundary layer which was triggered with a trip wire upstream of injection. The aerodynamic flow field is characterized in terms of profile Mach number distribution, and the associated heat transfer coefficients around the uncooled airfoil are presented. Both detailed and spanwise averaged results of film cooling effectiveness and heat transfer coefficients are shown on the suction side, which indicate considerable influence of the state of the incoming boundary layer on the performance of a film cooling row. The influence of the mainstream flow condition on the film cooling behavior at constant blowing ratio is discussed for three chosen injection regimes.

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

2015 ◽  
Author(s):  
Stephen P. Lynch ◽  
Karen A. Thole
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Vortical Flow ◽  
Leakage Flow ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
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.

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

2003 ◽  
Vol 125 (4) ◽  
pp. 648-657 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
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.

Effect of Blade Tip Geometry on Tip Flow and Heat Transfer for a Blade in a Low Speed Cascade

2003 ◽  
Author(s):  
Vikrant Saxena ◽  
Hasan Nasir ◽  
Srinath V. Ekkad
Keyword(s):  
Heat Transfer ◽  
Wind Tunnel ◽  
Flow Direction ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Transfer Coefficients ◽  
Wake Effect ◽  
Low Speed ◽  
Heat Transfer Coefficients ◽  
Blade Tip

A comprehensive investigation of the effect of various tip sealing geometries is presented on the blade tip leakage flow and associated heat transfer of a scaled up HPT turbine blade in a low-speed wind tunnel facility. The linear cascade is made of four blades with the two corner blades acting as guides. The tip section of a HPT first stage rotor blade is used to fabricate the 2-D blade. The wind tunnel accommodates an 116° turn for the blade cascade. The mainstream Reynolds number based on the axial chord length at cascade exit is 4.83 × 105. The upstream wake effect is simulated with a spoked wheel wake generator placed upstream of the cascade. A turbulence grid placed even farther upstream generates the required free-stream turbulence of 4.8%. The center blade has a tip clearance gap of 1.5625% with respect to the blade span. Static pressure measurements are obtained on the blade surface and the shroud. The effect of crosswise trip strips to reduce leakage flow and associated heat transfer is investigated with strips placed along the leakage flow direction, against the leakage flow and along the chord. Cylindrical pin fins and pitch variation of strips over the tip surface are also investigated. Detailed heat transfer measurements are obtained using a steady state HSI-based liquid crystal technique. The effect of periodic unsteady wake effect is also investigated by varying the wake Strouhal number from 0. to 0.2, and to 0.4. Results show that the trip strips placed against the leakage flow produce the lowest heat transfer on the tips compared to all the other cases with a reduction between 10–15% compared to the plain tip. Results also show that the pitch of the strips has a small effect on the overall reduction. Cylindrical pins fins and strips along the leakage flow direction do not decrease the heat transfer coefficients and in some cases enhance the heat transfer coefficients by as much as 20%.

Effect of Blade Tip Geometry on Tip Flow and Heat Transfer for a Blade in a Low-Speed Cascade

2004 ◽  
Vol 126 (1) ◽  
pp. 130-138 ◽  
Author(s):  
Vikrant Saxena ◽  
Hasan Nasir ◽  
Srinath V. Ekkad
Keyword(s):  
Heat Transfer ◽  
Wind Tunnel ◽  
Flow Direction ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Transfer Coefficients ◽  
Wake Effect ◽  
Low Speed ◽  
Heat Transfer Coefficients ◽  
Blade Tip

A comprehensive investigation of the effect of various tip sealing geometries is presented on the blade tip leakage flow and associated heat transfer of a scaled up HPT turbine blade in a low-speed wind tunnel facility. The linear cascade is made of four blades with the two corner blades acting as guides. The tip section of a HPT first stage rotor blade is used to fabricate the two-dimensional blade. The wind tunnel accommodates an 116 deg turn for the blade cascade. The mainstream Reynolds number based on the axial chord length at cascade exit is 4.83×105. The upstream wake effect is simulated with a spoked wheel wake generator placed upstream of the cascade. A turbulence grid placed even farther upstream generates the required freestream turbulence of 4.8%. The center blade has a tip clearance gap of 1.5625% with respect to the blade span. Static pressure measurements are obtained on the blade surface and the shroud. The effect of crosswise trip strips to reduce leakage flow and associated heat transfer is investigated with strips placed along the leakage flow direction, against the leakage flow and along the chord. Cylindrical pin fins and pitch variation of strips over the tip surface are also investigated. Detailed heat transfer measurements are obtained using a steady-state HSI-based liquid crystal technique. The effect of periodic unsteady wake effect is also investigated by varying the wake Strouhal number from 0. to 0.2, and to 0.4. Results show that the trip strips placed against the leakage flow produce the lowest heat transfer on the tips compared to all the other cases with a reduction between 10–15% compared to the plain tip. Results also show that the pitch of the strips has a small effect on the overall reduction. Cylindrical pins fins and strips along the leakage flow direction do not decrease the heat transfer coefficients and in some cases enhance the heat transfer coefficients by as much as 20%.

Measured Adiabatic Effectiveness and Heat Transfer for Blowing From the Tip of a Turbine Blade

2005 ◽  
Vol 127 (2) ◽  
pp. 251-262 ◽  
Author(s):  
J. R. Christophel ◽  
E. Couch ◽  
K. A. Thole ◽  
F. J. Cunha
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Suction Side ◽  
Leakage Flow ◽  
Transfer Coefficients ◽  
Forward Region ◽  
Cooling Effectiveness ◽  
Combustion Gas ◽  
Blade Tip ◽  
Blade Alloy

The clearance gap between the tip of a turbine blade and the shroud has an inherent leakage flow from the pressure side to the suction side of the blade. This leakage flow of combustion gas and air mixtures leads to severe heat transfer rates on the blade tip of the high-pressure turbine. As the thermal load to the blade increases, blade alloy oxidation and erosion rates increase thereby adversely affecting component life. The subject of this paper is the cooling effectiveness levels and heat transfer coefficients that result from blowing through two holes placed in the forward region of a blade tip. These holes are referred to as dirt purge holes and are generally required for manufacturing purposes and expelling dirt from the coolant flow when operating in sandy environments. Experiments were performed in a linear blade cascade for two tip-gap heights over a range of blowing ratios. Results indicated that the cooling effectiveness was highly dependent on the tip-gap clearance with better cooling achieved at smaller clearances. Also, heat transfer was found to increase with blowing. In considering an overall benefit of cooling from the dirt purge blowing, a large benefit was realized for a smaller tip gap as compared with a larger tip gap.

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

1995 ◽  
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 ◽  
Film 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.

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