Heat Transfer and Film-Cooling Measurements on a Stator Vane With Fan-Shaped Cooling Holes

2005 ◽  
Author(s):  
W. Colban ◽  
A. Gratton ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Pressure Side ◽  
Guide Vanes ◽  
Single Row ◽  
Inlet Guide Vanes ◽  
Adiabatic Effectiveness ◽  
Lift Off

In a typical gas turbine engine, the gas exiting the combustor is significantly hotter than the melting temperature of the turbine components. The highest temperatures in an engine are typically seen by the turbine inlet guide vanes. One method used to cool the inlet guide vanes is film-cooling, which involves bleeding comparatively low-temperature, high-pressure air from the compressor and injecting it through an array of discrete holes on the vane surface. To predict the vane surface temperatures in the engine, it is necessary to measure the heat transfer coefficient and adiabatic film-cooling effectiveness on the vane surface. This study presents heat transfer coefficients and adiabatic effectiveness levels measured in a scaled-up, two-passage cascade with a contoured endwall. Heat transfer measurements indicated that the behavior of the boundary layer transition along the suction side of the vane showed sensitivity to the location of film-cooling injection, which was simulated through the use of a trip wire placed on the vane surface. Single row adiabatic effectiveness measurements without any upstream blowing showed jet lift-off was prevalent along the suction side of the airfoil. Single row adiabatic effectiveness measurements on the pressure side, also without upstream showerhead blowing, indicated jet lifted-off and then reattached to the surface in the concave region of the vane. In the presence of upstream showerhead blowing, the jet lift-off for the first pressure side row was reduced, increasing adiabatic effectiveness levels.

Heat Transfer and Film-Cooling Measurements on a Stator Vane With Fan-Shaped Cooling Holes

2005 ◽  
Vol 128 (1) ◽  
pp. 53-61 ◽  
Author(s):  
W. Colban ◽  
A. Gratton ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Pressure Side ◽  
Guide Vanes ◽  
Single Row ◽  
Inlet Guide Vanes ◽  
Adiabatic Effectiveness ◽  
Lift Off

In a typical gas turbine engine, the gas exiting the combustor is significantly hotter than the melting temperature of the turbine components. The highest temperatures in an engine are typically seen by the turbine inlet guide vanes. One method used to cool the inlet guide vanes is film cooling, which involves bleeding comparatively low-temperature, high-pressure air from the compressor and injecting it through an array of discrete holes on the vane surface. To predict the vane surface temperatures in the engine, it is necessary to measure the heat transfer coefficient and adiabatic film-cooling effectiveness on the vane surface. This study presents heat transfer coefficients and adiabatic effectiveness levels measured in a scaled-up, two-passage cascade with a contoured endwall. Heat transfer measurements indicated that the behavior of the boundary layer transition along the suction side of the vane showed sensitivity to the location of film-cooling injection, which was simulated through the use of a trip wire placed on the vane surface. Single-row adiabatic effectiveness measurements without any upstream blowing showed jet lift-off was prevalent along the suction side of the airfoil. Single-row adiabatic effectiveness measurements on the pressure side, also without upstream showerhead blowing, indicated jet lifted-off and then reattached to the surface in the concave region of the vane. In the presence of upstream showerhead blowing, the jet lift-off for the first pressure side row was reduced, increasing adiabatic effectiveness levels.

Numerical Investigation of Coolant-to-Mainstream Scaling Parameters With Film Cooling on Pressure and Suction Side of a Gas Turbine Blade

2017 ◽  
Author(s):  
Lingyu Zeng ◽  
Xueying Li ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Momentum Flux ◽  
Velocity Ratio ◽  
Density Ratio ◽  
Flux Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

Most experiments of blade film cooling are conducted with density ratio lower than that of turbine conditions. In order to accurately model the performance of film cooling under a high density ratio, choosing an appropriate coolant to mainstream scaling parameter is necessary. The effect of density ratio on film cooling effectiveness on the surface of a gas turbine twisted blade is investigated from a numerical point of view. One row of film holes are arranged in the pressure side and two rows in the suction side. All the film holes are cylindrical holes with a pitch to diameter ratio P/d = 8.4. The inclined angle is 30°on the pressure side and 34° on the suction side. The steady solutions are obtained by solving Reynolds-Averaged-Navier-Stokes equations with a finite volume method. The SST turbulence model coupled with γ-θ transition model is applied for the present simulations. A film cooling experiment of a turbine vane was done to validate the turbulence model. Four different density ratios (DR) from 0.97 to 2.5 are studied. To independently vary the blowing ratio (M), momentum flux ratio (I) and velocity ratio (VR) of the coolant to the mainstream, seven conditions (M varying from 0.25 to 1.6 on the pressure side and from 0.25 to 1.4 on the suction side) are simulated for each density ratio. The results indicate that the adiabatic effectiveness increases with the increase of density ratio for a certain blowing ratio or a certain momentum flux ratio. Both on the pressure side and suction side, none of the three parameters listed above can serve as a scaling parameter independent of density ratio in the full range. The velocity ratio provides a relative better collapse of the adiabatic effectiveness than M and I for larger VRs. A new parameter describing the performance of film cooling is introduced. The new parameter is found to be scaled with VR for nearly the whole range.

Evaluation of Superposition Predictions for Showerhead Film Cooling on a Vane

2014 ◽  
Author(s):  
Joshua B. Anderson ◽  
James R. Winka ◽  
David G. Bogard ◽  
Michael E. Crawford
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Edge Region ◽  
Current Investigation ◽  
Engine Design ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

The leading edge of a turbine vane is subject to some of the highest temperature loading within an engine, and an accurate understanding of leading edge film coolant behavior is essential for modern engine design. Although there have been many investigations of the adiabatic effectiveness for showerhead film cooling of a vane leading edge region, there have been no previous studies in which individual rows of the showerhead were tested with the explicit intent of validating superposition models. For the current investigation, a series of adiabatic effectiveness experiments were performed with a five-row and three-row showerhead. The experiments were repeated separately with each individual row of holes active. This allowed evaluation of superposition methods on both the suction side of the vane, which was moderately convex, and the pressure side of the vane, which was mildly concave. Superposition was found to accurately predict performance on the suction side of the vane at lower momentum flux ratios, but not at higher momentum flux ratios. On the pressure side of the vane the superposition predictions were consistently lower than measured values, with significant errors occurring at the higher momentum flux ratios. Reasons for the under-prediction by superposition analysis are presented.

Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 1—Experimental Technique and Results

1992 ◽  
Vol 114 (4) ◽  
pp. 734-740 ◽  
Author(s):  
S. P. Harasgama ◽  
C. D. Burton
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Suction Side ◽  
Coolant Temperature ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

Heat transfer and aerodynamic measurements have been made on the endwalls of an annular cascade of turbine nozzle guide vanes in the presence of film cooling. The results indicate that high levels of cooling effectiveness can be achieved on the endwalls of turbine nozzle guide vanes (NGV). The NGV were operated at the correct engine nondimensional conditions of Reynolds number, Mach number, gas-to-wall temperature ratio, and gas-to-coolant density ratio. The results show that the secondary flow and horseshoe vortex act on the coolant, which is convected toward the suction side of the NG V endwall passage. Consequently the coolant does not quite reach the pressure side/casing trailing edge, leading to diminished cooling in this region. Increasing the blowing rate from 0.52 to 1.1 results in significant reductions in heat transfer to the endwall. Similar trends are evident when the coolant temperature is reduced. Measured heat transfer rates indicate that over most of the endwall region the film cooling reduces the Nusselt number by 50 to 75 percent.

Film Cooling Jet Injection Effect in Heat Transfer Coefficient Augmentation for the Pressure Side Cooling of Turbine Vane

2014 ◽  
Author(s):  
Hossein Nadali Najafabadi ◽  
Matts Karlsson ◽  
Mats Kinell ◽  
Esa Utriainen
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Time Resolved ◽  
Turbine Vane

Improving film cooling performance of turbine vanes and blades is often achieved through application of multiple arrays of cooling holes on the suction side, the showerhead region and the pressure side. This study investigates the pressure side cooling under the influence of single and multiple rows of cooling in the presence of a showerhead from a heat transfer coefficient augmentation perspective. Experiments are conducted on a prototype turbine vane working at engine representative conditions. Transient IR thermography is used to measure time-resolved surface temperature and the semi-infinite method is utilized to calculate the heat transfer coefficient on a low conductive material. Investigations are performed for cylindrical and fan-shaped holes covering blowing ratio 0.6 and 1.8 at density ratio of about unity. The freestream turbulence is approximately 5% close to the leading edge. The resulting heat transfer coefficient enhancement, the ratio of HTC with to that without film cooling, from different case scenarios have been compared to showerhead cooling only. Findings of the study highlight the importance of showerhead cooling to be used with additional row of cooling on the pressure side in order to reduce heat transfer coefficient enhancement. In addition, it is shown that extra rows of cooling will not significantly influence heat transfer augmentation, regardless of the cooling hole shape.

Film Cooling From a Single Row of Holes Oriented in Spanwise/Normal Planes

1997 ◽  
Vol 119 (4) ◽  
pp. 770-776 ◽  
Author(s):  
P. M. Ligrani ◽  
A. E. Ramsey
Keyword(s):  
Film Cooling ◽  
Test Surface ◽  
Spanwise Direction ◽  
Single Row ◽  
Local Maxima ◽  
Blowing Ratio ◽  
Streamwise Location ◽  
Structure Of Flow ◽  
Adiabatic Effectiveness ◽  
Lift Off

Experimental results are presented that describe the development and structure of flow downstream of a single row of film-cooling holes inclined at 30 deg from the test surface in spanwise/normal planes. With this configuration, holes are spaced 6d apart in the spanwise direction in a single row. Results are presented for a ratio of injectant density to free-stream density near 1.0, and injection blowing ratios from 0.5 to 1.5. Compared to results measured downstream of simple angle (streamwise) oriented holes, spanwise-averaged adiabatic effectiveness values are significantly higher for the same spanwise hole spacing, normalized streamwise location x/d, and blowing ratio m when m = 1.0 and 1.5 for x/d < 80. The injectant from the spanwise/normal holes is also less likely to lift off of the test surface than injectant from simple angle holes. This is because lateral components of momentum keep higher concentrations of injectant in closer proximity to the surface. As a result, local adiabatic effectiveness values show significantly greater spanwise variations and higher local maxima at locations immediately downstream of the holes. Spanwise-averaged iso-energetic Stanton number ratios range between 1.07 and 1.26, which are significantly higher than values measured downstream of two other injection configurations (one of which is simple angle, streamwise holes) when compared at the same x/d and blowing ratio.

Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 1 — Experimental Technique and Results

1991 ◽  
Author(s):  
S. P. Harasgama ◽  
C. D. Burton
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Suction Side ◽  
Coolant Temperature ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

Heat transfer and aerodynamic measurements have been made on the endwalls of an annular cascade of turbine nozzle guide vanes in the presence of film cooling. The results indicate that high levels of cooling effectiveness can be achieved on the endwalls of turbine nozzle guide vanes (NGV). The NGV were operated at the correct engine non-dimensional conditions of Reynolds number, Mach number, gas-to-wall temperature ratio and gas-to-coolant density ratio. The results show that the secondary flow and horse-shoe vortex act on the coolant which is converted towards the suction side of the NGV endwall passage. Consequently the coolant does not quite reach the pressure side/casing trailing edge, leading to diminished cooling in this region. Increasing the blowing rate from 0.52 to 1.1 results in significant reductions in heat transfer to the endwall. Similar trends are evident when the coolant temperature is reduced. Measured heat transfer rates indicate that over most of the endwall region the film cooling reduces the Nusselt number by 50% to 75%.

Effect of Incidence on Wall Heating Rates and Aerodynamics on a Film-Cooled Transonic Turbine Blade

1991 ◽  
Vol 113 (3) ◽  
pp. 493-501 ◽  
Author(s):  
C. Camci ◽  
T. Arts
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Convective Heat Transfer ◽  
Stagnation Point ◽  
Short Duration ◽  
Convective Heat ◽  
Film Cooling ◽  
Free Stream ◽  
Suction Side ◽  
Pressure Side

This study investigates the influence of incidence on convective heat transfer to highly curved surfaces of a film-cooled turbine rotor blade. A computational study of free-stream inviscid aerodynamics without cooling at various incidences is followed by well-documented measured heat transfer data sets. The heat transfer experiments are discussed for cases with and without film cooling, performed under realistic gas turbine flow conditions in the short-duration heat transfer facility of the von Karman Institute for Fluid Dynamics. The precise location of the stagnation point and the iso-Mach number contours in the passage for each incidence (−10, 0, 10, +15 deg) are presented for a nominal exit Mach number of 0.94. The free-stream mass flow rate was kept constant for each experiment at different incidence levels. Three rows of compound angled discrete cooling holes are located near the leading edge in a showerhead configuration. Two rows of staggered discrete cooling holes are located on the suction side and a single row of cooling holes is located on the pressure side. The short-duration measurements of quantitative wall heat fluxes on nearly isothermal blade surfaces both in the presence and absence of coolant ejection are presented. The study indicated that the change of the position of the stagnation point strongly altered the aerodynamic behavior and convective heat transfer to the blade in approximately the first 30 percent of both the pressure side and the suction side in the presence and absence of film cooling. The immediate vicinity of the stagnation point was not significantly affected by changing incidence without cooling. Transitional behavior both on the suction surface and on the pressure surface was significantly influenced by the changes in approaching flow direction. Flow separation associated with incidence variations was also observed. Extremely low levels of the convective heat transfer coefficients were experienced near the regions where small separation bubbles are located.

Effect of Incidence on Wall Heating Rates and Aerodynamics on a Film Cooled Transonic Turbine Blade

1990 ◽  
Author(s):  
Cengiz Camci ◽  
Tony Arts
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Convective Heat Transfer ◽  
Stagnation Point ◽  
Short Duration ◽  
Convective Heat ◽  
Film Cooling ◽  
Free Stream ◽  
Suction Side ◽  
Pressure Side

This study investigates the influence of incidence on convective heat transfer to highly curved surfaces of a film cooled turbine rotor blade. A computational study of free stream inviscid aerodynamics without cooling at various incidences is followed by well documented measured heat transfer data sets. The heat transfer experiments are discussed for cases with and without film cooling, performed under realistic gas turbine flow conditions in the short duration heat transfer facility of the von Karman Institute for Fluid Dynamics. The precise location of the stagnation point and the iso-Mach number contours in the passage for each incidence (−10°, 0°, 10°, +10°) are presented for a nominal exit Mach number of 0.94. The free stream mass flow rate was kept constant for each experiment at different incidence levels. Three rows of compound angled discrete cooling holes are located near the leading edge in a shower-head configuration. Two rows of staggered discrete cooling holes are located on the suction side and a single row of cooling holes is located on the pressure side. The short duration measurements of quantitative wall heat fluxes on nearly isothermal blade surfaces both in the presence and absence of coolant ejection are presented. The study indicated that the change of the position of the stagnation point strongly altered the aerodynamic behaviour and convective heat transfer to the blade in approximately the first 30 % of both the pressure side and the suction side in the presence and absence of film cooling. The immediate vicinity of the stagnation point was not significantly affected by changing incidence without cooling. Transitional behaviour both on the suction surface and on the pressure surface was significantly influenced by the changes in approching flow direction. Flow separation associated with incidence variations was also observed. Extremely low levels of convective heat transfer coefficients were experienced near the regions where small separation bubbles are located.

Turbine Blade Tip Film Cooling With Blade Rotation: Part I — Tip and Pressure Side Coolant Injection

2015 ◽  
Author(s):  
Onieluan Tamunobere ◽  
Sumanta Acharya
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
High Heat ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
High Heat Transfer ◽  
Blade Tip

This is the first in a two-part series of an experimental film cooling study conducted on the tip of a turbine blade with a blade rotation speed of 1200 RPM. In this part of the study, the coolant is injected from the blade tip and pressure side (PS) holes, and the effect of the blowing ratio on the heat transfer coefficient and film cooling effectiveness of the blade tip is investigated. The blade has a tip clearance of 1.7% of the blade span and consists of a cut back squealer rim, two cylindrical tip holes and six shaped pressure side holes. The stator-rotor-stator test section is housed in a closed loop wind tunnel that allows for the performance of transient heat transfer tests. Measurements of the heat transfer coefficient and film cooling effectiveness are done on the blade tip using liquid crystal thermography. These measurements are reported for the no coolant case and for blowing ratios of 1.0, 1.5, 2.0, 3.0 and 4.0. The heat transfer results for the no coolant injection show a region of high heat transfer on the blade tip near the blade leading edge region as the incident flow impinges on that region. This region of high heat transfer extends and stretches on the tip as more coolant is introduced through the tip holes at higher blowing ratios. The cooling results show that increasing the blowing ratio increases the film cooling effectiveness. The tip film cooling profile is such that the tip coolant is pushed towards the blade suction side thereby providing better coverage in that region. The shift in coolant flow profile towards the blade suction side as opposed to the pressure side in stationary studies can primarily be attributed to the effects of the blade relative motion.

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