scholarly journals Showerhead Film Cooling Performance of a Turbine Vane at High Freestream Turbulence in a Transonic Cascade

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Shakeel Nasir ◽  
Trey Bolchoz ◽  
Wing-Fai Ng ◽  
Luzeng J. Zhang ◽  
Hee Koo Moon ◽  
...  
Keyword(s):  
Reynolds Number ◽  
Nusselt Number ◽  
Mach Number ◽  
Film Cooling ◽  
Boundary Layer Transition ◽  
Suction Surface ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

This paper experimentally investigates the effect of blowing ratio and exit Reynolds number/Mach number on the film cooling performance of a showerhead film cooled first stage turbine vane. The vane midspan was instrumented with single-sided platinum thin film gauges to experimentally characterize the Nusselt number and film cooling effectiveness distributions over the surface. The vane was arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. Three different exit Mach numbers of Mex = 0.57, 0.76 and 1.0—corresponding to exit Reynolds numbers based on vane chord of 9.7 × 105, 1.1 × 106 and 1.5 × 106, respectively—were tested with an inlet free stream turbulence intensity (Tu) of 16% and an integral length scale normalized by vane pitch (Λx/P) of 0.23. A showerhead cooling scheme with five rows of cooling holes was tested at blowing ratios of BR = 0, 1.5, 2.0, and 2.5 and a density ratio of DR = 1.3. Nusselt number and adiabatic film cooling effectiveness distributions were presented on the vane surface over a range of s/C = −0.58 on the pressure side to s/C = 0.72 on the suction side of the vane. The primary effects of coolant injection were to augment the Nusselt number and reduce the adiabatic wall temperature downstream of the injection on the vane surface as compared to no film injection case (BR = 0) at all exit Mach number conditions. In general, an increase in blowing ratio (BR = 1.5 to 2.5) showed noticeable Nusselt number augmentation on pressure surface as compared to suction surface at exit Mach 0.57 and 0.75; however, the Nusselt number augmentation for these blowing ratios was found to be negligible on the vane surface for exit Mach 1.0 case. At exit Mach 1.0, an increase in blowing ratio (BR = 1.5 to 2.5) was observed to have an adverse effect on the adiabatic effectiveness on the pressure surface but had negligible effect on suction surface. The effectiveness trend on the suction surface was also found to be influenced by a favorable pressure gradient due to Mach number and boundary layer transition in the region s/C = 0.28 to s/C = 0.45 at all blowing ratio and exit Mach number conditions. An increase in Reynolds number from exit Mach 0.76 to 1.0 increased heat transfer levels on the vane surface at all blowing ratio conditions. A large increase in Reynolds number adversely affected adiabatic effectiveness on the pressure surface at all blowing ratio conditions. On the suction surface, a large increase in Reynolds number also affected adiabatic effectiveness in the favorable pressure gradient and boundary layer transition region.

scholarly journals Showerhead Film Cooling Performance of a Turbine Vane at High Freestream Turbulence in a Transonic Cascade

2008 ◽  
Author(s):  
S. Nasir ◽  
T. Bolchoz ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon ◽  
...  
Keyword(s):  
Reynolds Number ◽  
Nusselt Number ◽  
Mach Number ◽  
Film Cooling ◽  
Boundary Layer Transition ◽  
Suction Surface ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

This paper experimentally investigates the effect of blowing ratio and exit Reynolds number/Mach number on the film cooling performance of a showerhead film cooled first stage turbine vane. The vane midspan was instrumented with single-sided platinum thin film gauges to experimentally characterize the Nusselt number and film cooling effectiveness distributions over the surface. The vane was arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. Three different exit Mach numbers of Mex = 0.57, 0.76 and 1.0—corresponding to exit Reynolds numbers based on vane chord of 9.7 × 105, 1.1 × 106 and 1.5 × 106, respectively—were tested with an inlet free stream turbulence intensity (Tu) of 16% and an integral length scale normalized by vane pitch (Λx/P) of 0.23. A showerhead cooling scheme with five rows of cooling holes was tested at blowing ratios of BR = 0, 1.5, 2.0, and 2.5 and a density ratio of DR = 1.3. Nusselt number and adiabatic film cooling effectiveness distributions were presented on the vane surface over a range of s/C = −0.58 on the pressure side to s/C = 0.72 on the suction side of the vane. The primary effects of coolant injection were to augment the Nusselt number and reduce the adiabatic wall temperature downstream of the injection on the vane surface as compared to no film injection case (BR = 0) at all exit Mach number conditions. In general, an increase in blowing ratio (BR = 1.5 to 2.5) showed noticeable Nusselt number augmentation on pressure surface as compared to suction surface at exit Mach 0.57 and 0.75; however, the Nusselt number augmentation for these blowing ratios was found to be negligible on the vane surface for exit Mach 1.0 case. At exit Mach 1.0, an increase in blowing ratio (BR = 1.5 to 2.5) was observed to have an adverse effect on the adiabatic effectiveness on the pressure surface but had negligible effect on suction surface. The effectiveness trend on the suction surface was also found to be influenced by a favorable pressure gradient due to Mach number and boundary layer transition in the region s/C = 0.28 to s/C = 0.45 at all blowing ratio and exit Mach number conditions. An increase in Reynolds number from exit Mach 0.76 to 1.0 increased heat transfer levels on the vane surface at all blowing ratio conditions. A large increase in Reynolds number adversely affected adiabatic effectiveness on the pressure surface at all blowing ratio conditions. On the suction surface, a large increase in Reynolds number also affected adiabatic effectiveness in the favorable pressure gradient and boundary layer transition region.

Letterbox Trailing Edge Heat Transfer: Effects of Blowing Rate, Reynolds Number, and External Turbulence on Heat Transfer and Film Cooling Effectiveness

2009 ◽  
Vol 132 (1) ◽  
Author(s):  
N. J. Fiala ◽  
I. Jaswal ◽  
F. E. Ames
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Pressure Drop ◽  
Flow Rate ◽  
Film Cooling ◽  
Suction Surface ◽  
Trailing Edge ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

Heat transfer and film cooling distributions have been acquired for a vane trailing edge with letterbox partitions. Additionally, pressure drop data have been experimentally determined across a pin fin array and a trailing edge slot with letterbox partitions. The pressure drop across the array and letterbox trailing edge arrangement was measurably higher than for the gill slot geometry. Experimental data for the partitions and the inner suction surface region downstream from the slot have been acquired over a four-to-one range in vane exit condition Reynolds number (500,000, 1,000,000, and 2,000,000), with low (0.7%), grid (8.5%), and aerocombustor (13.5%) turbulence conditions. At these conditions, both heat transfer and adiabatic film cooling distributions have been documented over a range of blowing ratios (0.47≤M≤1.9). Heat transfer distributions on the inner suction surface downstream from the slot ejection were found to be dependent on both ejection flow rate and external conditions. Heat transfer on the partition side surfaces correlated with both exit Reynolds number and blowing ratio. Heat transfer on partition top surfaces largely correlated with exit Reynolds number but blowing ratio had a small effect at higher values. Generally, adiabatic film cooling levels on the inner suction surface are high but decrease near the trailing edge and provide some protection for the trailing edge. Adiabatic effectiveness levels on the partitions correlate with blowing ratio. On the partition sides adiabatic effectiveness is highest at low blowing ratios and decreases with increasing flow rate. On the partition tops adiabatic effectiveness increases with increasing blowing ratio but never exceeds the level on the sides. The present paper, together with a companion paper that documents letterbox trailing edge aerodynamics, is intended to provide engineers with the heat transfer and aerodynamic loss information needed to develop and compare competing trailing edge designs.

Letterbox Trailing Edge Heat Transfer: Effects of Blowing Rate, Reynolds Number, and External Turbulence on Heat Transfer and Film Cooling Effectiveness

2008 ◽  
Author(s):  
N. J. Fiala ◽  
I. Jaswal ◽  
F. E. Ames
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Pressure Drop ◽  
Flow Rate ◽  
Film Cooling ◽  
Suction Surface ◽  
Trailing Edge ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

Heat transfer and film cooling distributions have been acquired for a vane trailing edge with letterbox partitions. Additionally, pressure drop data have been experimentally determined across a pin fin array and a trailing edge slot with letterbox partitions. The pressure drop across the array and letterbox trailing edge arrangement was measurably higher than for the gill slot geometry. Experimental data for the partitions and the inner suction surface region downstream from the slot have been acquired over a four to one range in vane exit condition Reynolds number (500,000, 1,000,000 and 2,000,000), with low (0.7%), grid (8.5%), and aero-combustor (13.5%) turbulence conditions. At these conditions, both heat transfer and adiabatic film cooling distributions have been documented over a range of blowing ratios (0.47 ≤ M ≤ 1.9). Heat transfer distributions on the inner suction surface downstream from the slot ejection were found to be dependent on both ejection flow rate and external conditions. Heat transfer on the partition side surfaces correlated on both exit Reynolds number and blowing ratio. Heat transfer on partition top surfaces largely correlated on exit Reynolds number but blowing ratio had a small effect at higher values. Generally, adiabatic film cooling levels on the inner suction surface are high but decrease near the trailing edge and provide some protection for the trailing edge. Adiabatic effectiveness levels on the partitions correlate with blowing ratio. On the partition sides adiabatic effectiveness is highest at low blowing ratios and decreases with increasing flow rate. On the partition tops adiabatic effectiveness increases with increasing blowing ratio but never exceeds the level on the sides. The present paper, together with a companion paper which documents letterbox trailing edge aerodynamics, is intended to provide engineers with the heat transfer and aerodynamic loss information needed to develop and compare competing trailing edge designs.

scholarly journals Performance of a Showerhead and Shaped Hole Film Cooled Vane at High Freestream Turbulence and Transonic Conditions

2011 ◽  
Author(s):  
A. Newman ◽  
S. Xue ◽  
W. Ng ◽  
H. K. Moon ◽  
L. Zhang
Keyword(s):  
Reynolds Number ◽  
Turbulence Intensity ◽  
Film Cooling ◽  
Reynolds Numbers ◽  
Boundary Layer Transition ◽  
Length Scale ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Shaped Hole

An experimental study was performed to measure surface Nusselt number and film cooling effectiveness on a film cooled first stage nozzle guide vane using a transient thin film gauge (TFG) technique. The information presented attempts to further characterize the performance of shaped hole film cooling by taking measurements on a row of shaped holes downstream of leading edge showerhead injection on both the pressure and suction surfaces (hereafter PS and SS) of a 1st stage NGV. Tests were performed at engine representative Mach and Reynolds numbers and high inlet turbulence intensity and large length scale at the Virginia Tech Transonic Cascade facility. Three exit Mach/Reynolds number conditions were tested: 1.0/1,400,000; 0.85/1,150,000; and 0.60/850,000 where Reynolds number is based on exit conditions and vane chord. At Mach/Reynolds numbers of 1.0/1,450,000 and 0.85/1,150,000 three blowing ratio conditions were tested: BR = 1.0, 1.5, and 2.0. At a Mach/Reynolds number of 0.60/850,000, two blowing ratio conditions were tested: BR = 1.5 and 2.0. All tests were performed at inlet turbulence intensity of 12% and length scale normalized by the cascade pitch of 0.28. Film cooling effectiveness and heat transfer results compared well with previously published data, showing a marked effectiveness improvement (up to 2.5x) over the showerhead only NGV and agreement with published showerhead-shaped hole data. Net heat flux reduction was shown to increase substantially (average 2.6x) with the addition of shaped holes, with an increase (average 1.6x) in required coolant mass flow. Boundary layer transition location was shown to be within a consistent region on the suction side regardless of blowing ratio and exit Mach number.

Showerhead Film Cooling Performance of a Transonic Turbine Vane at High Freestream Turbulence (Tu = 16%): 3-D CFD and Comparison With Experiment

2008 ◽  
Author(s):  
Hong Wu ◽  
S. Nasir ◽  
W. F. Ng ◽  
H. K. Moon
Keyword(s):  
Nusselt Number ◽  
Mach Number ◽  
Film Cooling ◽  
Experimental Results ◽  
Suction Surface ◽  
Adiabatic Wall ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Favorable Pressure Gradient ◽  
Turbine Vane

The main objective of the study reported here is to use 3-D CFD to calculate and explain adiabatic film cooling effectiveness and Nusselt number distributions downstream of the showerhead film cooling rows of a turbine vane at high freestream turbulence and realistic exit Reynolds number/Mach number condition. The paper discusses a three-simulations technique to calculate vane surface recovery temperature, adiabatic wall temperature, and surface Nusselt number to completely characterize film cooling performance in a high speed flow. The RANS based ν2-f turbulence model, originally suggested by Durbin [1], is used in all numerical predictions. The vane midspan numerical calculations are compared with the experimental results obtained with the showerhead film cooled vane instrumented with single-sided platinum thin film gauges at the midspan and arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. Exit Mach number of Mex = 0.76—corresponding to exit Reynolds numbers based on vane chord of 1.1 × 106—was tested with an inlet free stream turbulence intensity (Tu) of 16% and an integral length scale normalized by vane pitch (Λx/P) of 0.23. A showerhead cooling scheme with five rows of cooling holes was tested at blowing ratios of BR = 0 and 1.5, and a density ratio of DR = 1.3. CFD predictions performed with experiment-matched boundary conditions show an overall good trend agreement with experimental adiabatic film cooling effectiveness and Nusselt number distributions downstream of the showerhead film cooling rows of the vane. For the experimental data, the primary effects of coolant injection are to augment Nusselt number and reduce adiabatic wall temperature downstream of the injection on the vane surface as compared to no film injection case (BR = 0) at Mex = 0.76. Similar to experimental results, the adiabatic film cooling effectiveness prediction on the suction surface at BR = 1.5 is found to be influenced by favorable pressure gradient due to Mach number through changes in local adiabatic wall and recovery temperature. The Nusselt number prediction on the suction surface shows a peak and a valley downstream of the film cooling rows in a favorable pressure gradient region for both tested blowing ratio conditions. This trend is also observed in the experimental results.

scholarly journals Turbine Nozzle Film Cooling Study Using the Pressure Sensitive Paint (PSP) Technique

1999 ◽  
Author(s):  
Luzeng Zhang ◽  
Michael Baltz ◽  
Ram Pudupatty ◽  
Michael Fox
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
High Speed ◽  
Surface Film ◽  
Pressure Sensitive Paint ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive

The use of pressure sensitive paint (PSP) to measure film cooling effectiveness on a turbine nozzle surface was demonstrated in a high speed wind tunnel. Film cooling effectiveness was measured from a single row of holes located on a turbine vane suction surface with a shaped exit. Nitrogen gas was used to simulate film cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Three blowing ratios were studied for each of the five freestream conditions: a reference condition, a reduced and an increased Reynolds number condition, and a reduced and an increased Mach number condition. The freestream turbulence intensity was kept at 12.0% for all the tests. The PSP was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the airfoil surface. The film effectiveness increased with blowing ratio for all the freestream conditions. The effects of secondary flow and freestream Mach number and Reynolds number on turbine nozzle suction surface film cooling are also discussed.

Experimental Investigation and Optimal Design on the Film Cooling Performance of Fan-Shaped Hole With Vortex Generator Fed by Crossflow

2021 ◽  
Author(s):  
Jie Wang ◽  
Chao Zhang ◽  
Xuebin Liu ◽  
Liming Song ◽  
Jun Li ◽  
...  
Keyword(s):  
Reynolds Number ◽  
Film Cooling ◽  
Vortex Generator ◽  
Optimized Design ◽  
Combined Model ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Shaped Hole

Abstract Aiming at investigating the effects of crossflow and vortex generator on film cooling characteristics of fan-shaped hole, the film cooling performance was measured experimentally by infrared camera. The blowing ratio is fixed at 0.5 and 1.5. The Reynolds number of the mainstream based on the hole diameter remains at 7000 and the inlet Reynolds number of crossflow is 40000. The experimental results show that the film cooling performance becomes better when the blowing ratio increases from 0.5 to 1.5 for each model, and the film cooling performance becomes worse under the influence of crossflow. When the blowing ratio is 1.5, the area-averaged film cooling effectiveness of the fan-shaped hole model with vortex generator decreases by 16.6% because of the influence of crossflow. The combined model always performs better compared with the model without vortex generator under all working conditions. When the blowing ratio becomes 1.5, under the influence of crossflow, the area-averaged film cooling effectiveness of the combined model could increase by 14.8%, compared with the model without vortex generator. To further improve the film cooling performance, the global optimization algorithm based on the Kriging method and the CFD technology are coupled to optimize the combined model under crossflow condition at the high blowing ratio, and the optimized design is verified by experiments. The experimental results show that the area-averaged film cooling effectiveness of the optimized design increases by 17.8% compared with the reference model.

Effectiveness of Normal and Angled Slot Cooling

2004 ◽  
Author(s):  
A. C. Smith ◽  
J. H. Hatchett ◽  
A. C. Nix ◽  
W. F. Ng ◽  
K. A. Thole ◽  
...  
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Fluid Dynamic ◽  
Current Experiment ◽  
Cfd Simulations ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Mach Numbers ◽  
Lift Off

An experimental and numerical investigation was conducted to determine the film cooling effectiveness of a normal slot and angled slot under realistic engine Mach number conditions. Freestream Mach numbers of 0.65 and 1.3 were tested. For the normal slot, hot gas ingestion into the slot was observed at low blowing ratios (M < 0.25). At high blowing ratios (M > 0.6) the cooling film was observed to “lift off” from the surface. For the 30° angled slot, the data was found to collapse using the blowing ratio as a scaling parameter. Results from the current experiment were compared with the subsonic data previously published. For the angle slot, at supersonic freestream Mach number, the current experiment shows that at the same x/Ms, the film-cooling effectiveness increases by as much as 25% as compared to the subsonic case. The results of the experiment also show that at the same x/Ms, the film cooling effectiveness of the angle slot is considerably higher than the normal slot, at both subsonic and supersonic Mach numbers. The flow physics for the slot tests considered here are also described with computational fluid dynamic (CFD) simulations in the subsonic and supersonic regimes.

scholarly journals Film Cooling on a Gas Turbine Rotor Blade

1991 ◽  
Author(s):  
Kenichiro Takeishi ◽  
Sunao Aoki ◽  
Tomohiko Sato ◽  
Keizo Tsukagoshi
Keyword(s):  
Mass Transfer ◽  
Film Cooling ◽  
Rotor Blade ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Rotor Blade ◽  
Pressure Surface ◽  
Rotating Blade

The film cooling effectiveness on a low-speed stationary cascade and the rotating blade has been measured by using a heat-mass transfer analogy. The film cooling effectiveness on the suction surface of the rotating blade fits well with that on the stationary blade, but a low level of effectiveness appears on the pressure surface of the rotating blade. In this paper, typical film cooling data will be presented and film cooling on a rotating blade is discussed.

Transonic Aerodynamic Losses Due to Turbine Airfoil, Suction Surface Film Cooling

1999 ◽  
Vol 122 (2) ◽  
pp. 317-326 ◽  
Author(s):  
D. J. Jackson ◽  
K. L. Lee ◽  
P. M. Ligrani ◽  
P. D. Johnson
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
Surface Film ◽  
Suction Surface ◽  
Pressure Losses ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Mixing Losses ◽  
Aerodynamic Losses

The effects of suction surface film cooling on aerodynamic losses are investigated using an experimental apparatus designed especially for this purpose. A symmetric airfoil with the same transonic Mach number distribution on both sides is employed. Mach numbers range from 0.4 to 1.24 and match values on the suction surface of airfoils from operating aeroengines. Film cooling holes are located on one side of the airfoil near the passage throat where the free-stream Mach number is nominally 1.07. Round cylindrical and conical diffused film cooling hole configurations are investigated with density ratios from 0.8 to 1.3 over a range of blowing ratios, momentum flux ratios, and Mach number ratios. Also included are discharge coefficients, local and integrated total pressure losses, downstream kinetic energy distributions, Mach number profiles, and a correlation for integral aerodynamic losses as they depend upon film cooling parameters. The contributions of mixing and shock waves to total pressure losses are separated and quantified. These results show that losses due to shock waves vary with blowing ratio as shock wave strength changes. Aerodynamic loss magnitudes due to mixing vary significantly with film cooling hole geometry, blowing ratio, Mach number ratio, and (in some situations) density ratio. Integrated mixing losses from round cylindrical holes are three times higher than from conical diffused holes, when compared at the same blowing ratio. Such differences depend upon mixing losses just downstream of the airfoil, as well as turbulent diffusion of streamwise momentum normal to the airfoil symmetry plane. [S0889-504X(00)02202-9]

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