The Effect of Compound Angle on Nozzle Suction Side Film Cooling

2012 ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Film Cooling ◽  
Flow Direction ◽  
Suction Side ◽  
Operating Conditions ◽  
The Other ◽  
Counter Flow ◽  
Film Cooling Effectiveness ◽  
Test Models ◽  
Angle Effect ◽  
Compound Angle

The film cooling hole compound angle effect on nozzle suction side film cooling effectiveness was experimentally investigated using a single row of shaped holes. Engine operating conditions were simulated in a scaled warm cascade, which was built based on industrial gas turbine nozzle vanes. Local film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness can be obtained by the mass transfer analogy. Three separate nozzle test models were fabricated, which have the same cooling supply plenum configurations. The baseline configuration has a row of shaped holes on the suction surface without a compound angle. The other two test models have same size film holes at the same location, but one with a compound angle in co-flow and the other in counter-flow direction to the cooling supply. Four cooling mass flow ratios (MFR) were studied for each of the nozzle test models and two-dimensional film effectiveness distributions were measured. Then the film effectiveness distributions were spanwise averaged for comparison. The compound angle injections prevent jets lift-off for higher MFR cases, but for a lower MFR, the baseline configuration results in higher film effectiveness. The co-flow or counter-flow compound angle results in film shifting to upper or lower endwall, but the effect of the direction of the compound angle is less significant compared to the compound angle effect.

The Effect of Wall Thickness on Nozzle Suction Side Film Cooling

2008 ◽  
Author(s):  
Luzeng Zhang ◽  
Hee Koo Moon
Keyword(s):  
Wall Thickness ◽  
Film Cooling ◽  
Suction Side ◽  
Operating Conditions ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
Test Models ◽  
Lift Off ◽  
Gas Turbine Nozzle ◽  
Industrial Gas Turbine

The effect of nozzle airfoil wall thickness on nozzle suction side film cooling effectiveness was experimentally investigated. The engine operating conditions were simulated in a scaled warm cascade, which was built based on industrial gas turbine nozzle airfoils. Film effectiveness distribution on the suction surfaces was obtained using the pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Two nozzle test models were investigated, which had similar cooling supply plenums. One of them had a wall thickness of 0.47cm and the other 0.63 cm. The nozzle vanes had showhead and film cooling holes on two suction side locations. Six cooling mass flow ratios (MFR, blowing ratios) were studied for each of the nozzle test models and two-dimensional film effectiveness distributions were obtained. Then the effectiveness distributions were spanwise averaged for comparison. For both cases, the overall film effectiveness increased with the MFR or the blowing ratio. Comparing the two models, the thicker wall (or the larger L/D value) resulted in improved film effectiveness for the lower MFR’s. For the higher MFR, the better-developed jets due to the thicker wall (or the larger L/D) would cause a jet lift-off; therefore reduce the overall film effectiveness. The spanwise shifting pattern was also depended on the wall thickness.

The Effect of Compound Angle on Nozzle Pressure Side Film Cooling

2009 ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Film Cooling ◽  
Operating Conditions ◽  
The Other ◽  
Hole Injection ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Test Models ◽  
Nozzle Pressure ◽  
Shaped Hole ◽  
Compound Angle

The effect of film cooling hole compound angle on nozzle pressure side film cooling effectiveness was experimentally investigated using a single row of shaped hole injection. The engine operating conditions were simulated in a scaled warm cascade, which was built based on industrial gas turbine nozzle vanes. Local film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy. Three separate nozzle test models were fabricated, which have same cooling supply plenum configurations. One of them has a row of shaped hole on the pressure surface without a compound angle. The other two test models have same size film holes at the same location, but one with a 30-degree compound angle in co-flow and the other in counter-flow direction to the cooling supply. Four cooling mass flow ratios (MFR, blowing ratio) were studied for each of the nozzle test models and two-dimensional film effectiveness distributions were measured. Then the film effectiveness distributions were spanwise averaged for comparison. For all three cases, the overall film effectiveness increased with the MFR (or the blowing ratio), but not significantly. Film effectiveness by a compound angle injection is higher compared to those without a compound angle near the injection, further downstream the difference is insignificant.

Film Cooling Effectiveness of Transonic Turbine Vane Suction Side With Compound-Angle Shaped Hole Configuration

2015 ◽  
Author(s):  
Shang-Feng Yang ◽  
Je-Chin Han ◽  
Alexander MirzaMoghadam ◽  
Ardeshir Riahi
Keyword(s):  
Transonic Flow ◽  
Film Cooling ◽  
Surface Film ◽  
Density Ratio ◽  
Suction Side ◽  
Flow Loop ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Compound Angle

This paper studies the effect of transonic flow velocity on local film cooling effectiveness distribution of turbine vane suction side, experimentally. A conduction-free Pressure Sensitive Paint (PSP) method is used to determine the local film cooling effectiveness. Tests were performed in a five-vane annular cascade at Texas A&M Turbomachinery laboratory blow-down flow loop facility. The exit Mach numbers are controlled to be 0.7, 0.9, and 1.1, from subsonic to transonic flow conditions. Three foreign gases N2, CO2 and Argon/SF6 mixture are selected to study the effects of three coolant-to-mainstream density ratios, 1.0, 1.5, and 2.0 on film cooling. Four averaged coolant blowing ratios in the range, 0.7, 1.0, 1.3 and 1.6 are investigated. The test vane features 3 rows of radial-angle cylindrical holes around the leading edge, and 2 rows of compound-angle shaped holes on the suction side. Results suggest that the PSP technique is capable of producing clear and detailed film cooling effectiveness contours at transonic condition. The effects of coolant to mainstream blowing ratio, density ratio, and exit Mach number on the vane suction-surface film cooling distribution are obtained, and the consequence results are presented and explained in this investigation.

Practical Slot Configurations for Turbine Film Cooling Applications

2009 ◽  
Author(s):  
Joshua E. Bruce-Black ◽  
Frederick T. Davidson ◽  
David G. Bogard ◽  
David R. Johns
Keyword(s):  
Film Cooling ◽  
Flow Characteristics ◽  
Suction Side ◽  
Operating Conditions ◽  
Coolant Flow ◽  
Film Cooling Effectiveness ◽  
Diffusion Technique ◽  
Turbine Vane ◽  
Structural Weakness ◽  
Turbine Component

Turbine component film cooling is most effective when using a continuous slot to introduce coolant to the surface. However, this is not practical due to the structural weakness that would be inherent with a continuous slot. In this study, several slot-like designs are investigated to establish the film cooling effectiveness. These slot configurations extended only a partial distance through the simulated turbine vane wall, and were fed with impinging cylindrical holes. The configurations were studied on the suction side of a scaled-up turbine vane. In this study varying slot widths, discrete and continuous slots, and diffusing the coolant flow within the slot prior to it being emitted onto the surface of the vane were investigated. Rows of discrete round and shaped holes were also tested for comparison with the slots. The study of varying slot geometries showed that decreasing the width of the slots led to a substantial increase in adiabatic effectiveness. An internal coolant diffusion technique showed promise by maintaining performance levels while potentially providing a design configuration that more readily meets structural demands in real world operating conditions. The coolant flow characteristics were also studied through the use of thermal profiles measurements. These thermal profiles showed significant mainstream ingestion on the top surface of the slot prior to the coolant emitting onto the surface of the vane.

Film Cooling Effectiveness Distribution on First-Stage Vane Endwall With and Without Leading-Edge Fillets: Part I—Effect of Leading Edge Geometry

2011 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Flow Direction ◽  
Leading Edge ◽  
Suction Side ◽  
Axial Direction ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Geometry ◽  
Blade Cascade

The paper is focused on the effect of leading edge airfoil geometry on endwall film cooling. Fillets placed at the junctions of the leading edge and the endwall are used in investigation. Three types of fillet profiles are tested, and the results are compared with baseline geometry without fillet. The design of the fillet is based on the suggestion by previous literature data indicating that sharp is effective in controlling the secondary flow. Three types of sharp slope fillet with the length to height ratio of 2.8, 1.2 and 0.5 are made using stereo lithography (SLA) and assessed in the experiment. Distributed with the approximately inviscid flow direction, four rows of compound angle laidback fan-shaped holes are arranged on the endwall to form full covered coolant film. The four rows of fanshaped holes are inclined 30 deg to the endwall surface and held an angle of 0, 30, 45 and 60 deg to axial direction respectively. The fanshaped holes have a lateral diffusion angle of 10 deg from the hole-centerline and a forward expansion angle of 10 deg to the endwall surface. The Reynolds number based on the axial chord and inlet velocity of the free-stream flow is 3.5*105, and the testing is done in a four-blade cascade with low Mach number condition (0.1 at the inlet) while the blowing ratio of the coolant through the discrete holes varies from 0.4 to 1.2. The film-cooling effectiveness distributions are obtained using the PSP (pressure sensitive paint) technique, by which the effect of different fillet geometry on passage induced flow and coolant is shown. The present paper compares the film cooling effectiveness distributions in a baseline blade cascade with three similar blades with different leading edge by adding fillets. The results show that with blowing ratio increasing, the film cooling effectiveness increases on the endwall. For specific blowing ratio, the effects of leading edge geometries could be illustrated as follows. The baseline geometry provides the best film cooling performance near leading edge pressure side. As for the leading edge suction side, the best leading edge geometry depends on the blowing ratio. The longfillet is the more effective in controlling horseshoe vortex at low blowing ratio, but for the high blowing ratio shortfillet and mediumfillet are better.

The Effects of Vane Showerhead Injection Angle and Film Compound Angle on Nozzle Endwall Cooling (Phantom Cooling)

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Nitrogen Gas ◽  
Injection Angle ◽  
Test Models ◽  
Cooling Hole ◽  
Cooling Effects ◽  
Film Cooling Holes ◽  
Compound Angle ◽  
Endwall Cooling

The effects of airfoil showerhead (SH) injection angle and film-cooling hole compound angle on nozzle endwall cooling (second order film-cooling effects, also called "phantom cooling") were experimentally investigated in a scaled linear cascade. The test cascade was built based on a typical industrial gas turbine nozzle vane. Endwall surface phantom cooling film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness can be obtained by the mass transfer analogy. Two separate nozzle test models were fabricated, which have the same number and size of film-cooling holes but different configurations. One had a SH angle of 45 deg and no compound angles on the pressure and suction side (SS) film holes. The other had a 30 deg SH angle and 30 deg compound angles on the pressure and SS film-cooling holes. Nitrogen gas (cooling air) was fed through nozzle vanes, and measurements were conducted on the endwall surface between the two airfoils where no direct film cooling was applied. Six cooling mass flow ratios (MFRs, blowing ratios) were studied, and local (phantom) film effectiveness distributions were measured. Film effectiveness distributions were pitchwise averaged for comparison. Phantom cooling on the endwall by the SS film injections was found to be insignificant, but phantom cooling on the endwall by the pressure side (PS) airfoil film injections noticeably helped the endwall cooling (phantom cooling) and was a strong function of the MFR. It was concluded that reducing the SH angle and introducing a compound angle on the PS injections would enhance the endwall surface phantom cooling, particularly for a higher MFR.

Influence of Coolant Density on Turbine Blade Film-Cooling With Compound-Angle Shaped Holes

2012 ◽  
Author(s):  
Kevin Liu ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Density Ratio ◽  
Flux Ratio ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Shaped Hole ◽  
Compound Angle

Adiabatic film-cooling effectiveness is examined systematically on a typical high pressure turbine blade by varying three critical flow parameters: coolant blowing ratio, coolant-to-mainstream density ratio, and freestream turbulence intensity. Three average coolant blowing ratios 1.0, 1.5, and 2.0; three coolant density ratios 1.0, 1.5, and 2.0; two turbulence intensities 4.2% and 10.5%, are chosen for this study. Conduction-free pressure sensitive paint (PSP) technique is used to measure film-cooling effectiveness. Three foreign gases — N2 for low density, CO2 for medium density, and a mixture of SF6 and Argon for high density are selected to study the effect of coolant density. The test blade features 45° compound-angle shaped holes on the suction side and pressure side, and 3 rows of 30° radial-angle cylindrical holes around the leading edge region. The inlet and the exit Mach number are 0.27 and 0.44, respectively. Reynolds number based on the exit velocity and blade axial chord length is 750,000. Results reveal that the PSP is a powerful technique capable of producing clear and detailed film effectiveness contours with diverse foreign gases. As blowing ratio exceeds the optimum value, it induces more mixing of coolant and mainstream. Thus film-cooling effectiveness reduces. Greater coolant-to-mainstream density ratio results in lower coolant-to-mainstream momentum and prevents coolant to lift-off; as a result, film-cooling increases. Higher freestream turbulence causes effectiveness to drop everywhere except in the region downstream of suction side. Results are also correlated with momentum flux ratio and compared with previous studies. It shows that compound shaped hole has the greatest optimum momentum flux ratio, and then followed by axial shaped hole, compound cylindrical hole, and axial cylindrical hole.

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

2008 ◽  
Vol 131 (1) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta P. Narzary ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Compound Angle

The film-cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint technique. Compound angle laidback fan-shaped holes are used to cool the blade surface with four rows on the pressure side and two rows on the suction side. The coolant injects to one side of the blade, either pressure side or suction side. The presence of wake due to the upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate progressing wakes. The effect of wakes is recorded at four phase locations along the pitchwise direction. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The inlet and exit Mach numbers are 0.27 and 0.44, respectively, resulting in a pressure ratio of 1.14. Five average blowing ratios ranging from 0.4 to 1.5 are tested. Results reveal that the tip-leakage vortices and endwall vortices sweep the coolant on the suction side to the midspan region. The compound angle laidback fan-shaped holes produce a good film coverage on the suction side except for the regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and the effectiveness level is low on the pressure surface. However, the pressure side acquires a relatively uniform film coverage with the multiple rows of cooling holes. The film-cooling effectiveness increases with the increasing average blowing ratio for either side of coolant ejection. The presence of stationary upstream wake results in lower film-cooling effectiveness on the blade surface. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film-cooling effectiveness, particularly at higher blowing ratios.

High Resolution Film Cooling Effectiveness Comparison of Axial and Compound Angle Holes on the Suction Side of a Turbine Vane

2006 ◽  
Author(s):  
Scot K. Waye ◽  
David G. Bogard
Keyword(s):  
High Resolution ◽  
Flat Plate ◽  
Film Cooling ◽  
Density Ratio ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Adiabatic Effectiveness ◽  
Compound Angle

Film cooling adiabatic effectiveness for axial and compound angle holes on the suction side of a simulated turbine vane was investigated to determine the relative performance of these configurations. The effect of the surface curvature was also evaluated by comparing to previous curvature studies and flat plate film cooling results. Experiments were conducted for varying coolant density ratio, mainstream turbulence levels, and hole spacing. Results from these measurements showed that for mild curvature, 2r/d ≈ 160, flat plate results are sufficient to predict the cooling effectiveness. Furthermore, the compound angle injection improves adiabatic effectiveness for higher blowing ratios, similar to previous studies using flat plate facilities.

Influence of Shock Wave on Turbine Vane Suction Side Film Cooling With Compound-Angle Shaped Holes

2011 ◽  
Author(s):  
K. Liu ◽  
D. P. Narzary ◽  
J. C. Han ◽  
A. V. Mirzamoghadam ◽  
A. Riahi
Keyword(s):  
Transonic Flow ◽  
Film Cooling ◽  
Flux Ratio ◽  
Suction Side ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Compound Angle

This paper studies the effect of shock wave on turbine vane suction side film cooling using a conduction-free Pressure Sensitive Paint (PSP) technique. Tests were performed in a five-vane annular cascade with a blow-down flow loop facility. The exit Mach numbers are controlled to be 0.7, 1.1, and 1.3, from subsonic to transonic flow conditions. Two foreign gases N2 and CO2 are selected to study the effects of two coolant-to-mainstream density ratios, 1.0 and 1.5, on film cooling. Four averaged coolant blowing ratios in the range, 0.4 to 1.6 are investigated. The test vane features 3 rows of radial-angle cylindrical holes around the leading edge, and 2 rows of compound-angle shaped holes on the suction side. Results suggest that the PSP is an accurate technique capable of producing clear and detailed film cooling effectiveness contours at transonic flow conditions. At lower blowing ratio, film cooling effectiveness decreases with increasing exit Mach number. On the other hand, an opposite trend is observed at high blowing ratio. In transonic flow, the rapid rise in pressure caused by shock benefits film-cooling by deflecting the coolant jet toward the vane surface at higher blowing ratio. Results show that denser coolant performs better, typically at higher blowing ratio in transonic flow. Results also show that the optimum momentum flux ratio decreases with density ratio at subsonic condition. In transonic flow, however, the trend is reversed and the peak effectiveness values plateau over a long range of momentum flux ratio.

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