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.

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

2007 ◽  
Author(s):  
Zhihong Gao ◽  
Diganta P. Narzary ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Steady Increase ◽  
Pressure Side ◽  
Wake Effect ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Compound Angle

The film cooling effectiveness on the surface of a high pressure turbine blade is measured using the Pressure Sensitive Paint (PSP). Four rows of fan-shaped, laid-back compound angled cooling holes are distributed on the pressure side while two such rows are provided on the suction side of the blade. The coolant is only injected to either the pressure side or suction side of the blade at five average blowing ratios from 0.4 to 1.5. Presence of wake due to 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 a progressing wake. Effect of wake is recorded at four phase locations with equal intervals along the pitch-wise direction. The free stream Reynolds number, based on the axial chord length and the exit velocity, is 750,000 and the inlet and the exit Mach numbers are 0.27 and 0.44, respectively, resulting in a blade pressure ratio of 1.14. Results reveal that the tip leakage vortices and endwall vortices sweep the coolant film on the suction side to the midspan region. The fan-shaped, laid-back compound angled holes produce good coolant film coverage on the suction side except for those regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and effectiveness level is low on the pressure surface. However, the pressure side acquires relatively uniform film coverage with the design of multiple rows of cooling holes. The presence of stationary upstream wake results in lower film cooling effectiveness on the blade surface. Variation of blowing ratio from 0.4 to 1.5 shows steady increase in effectiveness on the pressure side or the suction side for a given wake rod phases locations. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film cooling effectiveness particularly at higher blowing ratios.

Effect of Unsteady Wake on Film-Cooling Effectiveness Distribution on a Gas Turbine Blade With Compound Shaped Holes

2007 ◽  
Author(s):  
Diganta P. Narzary ◽  
Zhihong Gao ◽  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Intensity Level ◽  
Pressure Side ◽  
Wake Effect ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cylindrical Holes

The effect of fan-shaped, laid-back compound angled cooling holes placed along the span of a fully-cooled high pressure turbine blade in a 5-blade linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Four rows of shaped film cooling holes are provided on the pressure side while two such rows are provided on the suction side of the blade. Three rows of cylindrical holes are drilled at 30° to the surface on the leading edge to capture the effect of showerhead film coolant injection. The coolant is injected at four different average blowing ratios of 0.3, 0.6, 0.9 and 1.2. Presence of wake due to upstream vanes is studied by placing a periodic set of rods upstream of the test blade. The wake is generated using 4.8mm diameter rods. The wake rods can be clocked by changing their stationary positions in front of the test blade to simulate a progressing wake. Effect of wake is recorded at four phase locations with equal intervals. The free stream Reynolds number, based on the axial chord length and the exit velocity, is 750,000 and the inlet and the exit Mach numbers are 0.27 and 0.44, respectively resulting in a blade pressure ratio of 1.14. Turbulence intensity level at the cascade inlet is 6% with an integral length scale of around 5cm. Results show that the fan-shaped, laid-back compound angled holes produce uniform and wide coolant coverage on the suction side except for those regions affected by the passage and tip leakage vortices. The advantage of compound shaped hole design is seen from the higher effectiveness values on the suction side compared to that of the compound cylindrical holes. The presence of a stationary upstream wake can result in lower film cooling effectiveness on the blade surface. Variation of blowing ratio from 0.3 to 1.2 show more or less uniform increment in effectiveness increase on the pressure side, whereas on the suction side, the increment shows signs of saturation beyond M = 0.6.

Influence of Coolant Density on Turbine Blade Film-Cooling Using Pressure Sensitive Paint Technique

2011 ◽  
Vol 134 (3) ◽  
Author(s):  
Diganta P. Narzary ◽  
Kuo-Chun Liu ◽  
Akhilesh P. Rallabandi ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Pressure Sensitive Paint ◽  
Freestream Turbulence ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive

Adiabatic film-cooling effectiveness is examined on a high-pressure turbine blade by varying three critical engine parameters, viz., coolant blowing ratio, coolant-to-mainstream density ratio, and freestream turbulence intensity. Three average coolant blowing ratios (BR=1.2, 1.7, and 2.2 on the pressure side and BR=1.1, 1.4, and 1.8 on the suction side), three average coolant density ratios (DR=1.0, 1.5, and 2.5), and two average freestream turbulence intensities (Tu=4.2% and 10.5%) are considered. Conduction-free pressure sensitive paint (PSP) technique is adopted 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 two rows of cylindrical film-cooling holes on the suction side (45 deg compound), 4 rows on the pressure side (45 deg compound) and 3 around the leading edge (30 deg radial). The inlet and the exit Mach numbers are 0.24 and 0.44, respectively. The Reynolds number of the mainstream flow is 7.5×105 based on the exit velocity and blade chord length. Results suggest that the PSP is a powerful technique capable of producing clear and detailed film-effectiveness contours with diverse foreign gases. Large improvement on the pressure side and moderate improvement on the suction side effectiveness is witnessed when blowing ratio is raised from 1.2 to 1.7 and 1.1 to 1.4, respectively. No major improvement is seen thereafter with the downstream half of the suction side showing drop in effectiveness. The effect of increasing coolant density is to increase effectiveness everywhere on the pressure surface and suction surface except for the small region on the suction side, xss/Cx<0.2. Higher freestream turbulence causes effectiveness to drop everywhere except in the region downstream of the suction side where significant improvement in effectiveness is seen.

A Numerical Study on the Characteristics of How Heat and Cooling Transfer on the Leading Edge of a Film-Cooling Blade

2011 ◽  
Vol 383-390 ◽  
pp. 3963-3968
Author(s):  
Shao Hua Li ◽  
Li Mei Du ◽  
Wen Hua Dong ◽  
Ling Zhang
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Pressure Surface

In this paper, a numerical simulation was performed to investigate heat transferring characteristics on the leading edge of a blade with three rows of holes of film-cooling using Realizable k- model. Three rows of holes were located on the suction side leading edge stagnation line and the pressure surface. The difference of the cooling efficiency and the heat transfer of the three rows of holes on the suction side and pressure side were analyzed; the heat transfer and film cooling effectiveness distribution in the region of leading edge are expounded under different momentum rations.The results show that under the same condition, the cooling effectiveness on the pressure side is more obvious than the suction side, but the heat transfer is better on the suction side than the pressure side. The stronger momentum rations are more effective cooling than the heat transfer system.

Effect of Superposition on Spanwise Film-Cooling Effectiveness Distribution on a Gas Turbine Blade

2006 ◽  
Author(s):  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Axial Direction ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Full Coverage ◽  
Film Cooling Holes

The effect of film cooling holes placed along the span of high pressure turbine blade in a 5 bladed linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Four rows of film cooling holes are provided on the pressure side while two such rows are provided on the suction side of the blade. Around 22 cylindrical holes with a diameter of 0.65mm are drilled in each row at a compound angle of 45° to the blade span in the radial direction and at 45° in the axial direction. Film cooling effectiveness over the entire blade region is determined from full coverage film cooling with coolant blowing from all rows and from each individual row. The effect of superposition of film cooling effectiveness from each individual row is then compared with full coverage film cooling. The coolant is injected at four different average blowing ratios of 0.6, 0.9, 1.2 and 1.5. The free stream Reynolds number, based on the axial chord length and the exit velocity, is 750,000 and the inlet and the exit Mach numbers are 0.27 and 0.44, respectively resulting in a blade pressure ratio of 1.14. Turbulence intensity level at the cascade inlet is 6% with an integral length scale of around 5cm. Results show that the effectiveness magnitudes from superposition of effectiveness data from individual rows are comparable with that from full coverage film cooling. Varying blowing ratios can have a significant impact on film-cooling effectiveness distribution with a blowing ratio of 0.6 showing highest effectiveness immediately downstream of the holes.

Influence of Coolant Density on Turbine Blade Film-Cooling Using Pressure Sensitive Paint Technique

2010 ◽  
Author(s):  
Diganta P. Narzary ◽  
Kuo-Chun Liu ◽  
Akhilesh P. Rallabandi ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Pressure Sensitive Paint ◽  
Freestream Turbulence ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive

Adiabatic film-cooling effectiveness is examined on a high pressure turbine blade by varying three critical engine parameters, viz., coolant blowing ratio, coolant-to-mainstream density ratio and freestream turbulence intensity. Three average coolant blowing ratios (BR = 1.2, 1.7, and 2.2 on the pressure side and BR = 1.1, 1.4, and 1.8 on the suction side), three average coolant density ratios (DR = 1.0, 1.5, and 2.5), and two average freestream turbulence intensities (Tu = 4.2% and 10.5%) are considered. Conduction-free Pressure Sensitive Paint (PSP) technique is adopted 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 2 rows of cylindrical film-cooling holes on the suction side (45° compound), 4 rows on the pressure side (45° compound) and 3 around the leading edge (30° radial). The inlet and the exit Mach numbers are 0.24 and 0.44, respectively. Reynolds number of the mainstream flow is 7.5E105 based on the exit velocity and blade chord length. Results suggest that the PSP is a powerful technique capable of producing clear and detailed film effectiveness contours with diverse foreign gases. Large improvement on the pressure side and moderate improvement on the suction side effectiveness is witnessed when blowing ratio is raised from 1.2 to 1.7 and 1.1 to 1.4, respectively. No major improvement is seen thereafter with the downstream half of the suction side showing drop in effectiveness. The effect of increasing coolant density is to increase effectiveness everywhere on the pressure surface and suction surface except for the small region on the suction side, xss/Cx&lt;0.2. Higher freestream turbulence causes effectiveness to drop everywhere except in the region downstream of the suction side where significant improvement in effectiveness is seen.

3D Flow Prediction and Improvement of Holes Arrangement of a Film-Cooled Turbine Blade Using a Feature-Based Jet Model

2006 ◽  
Author(s):  
Andre´ Burdet ◽  
Reza S. Abhari
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Good Prediction ◽  
Pressure Side ◽  
3D Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Numerical Predictions ◽  
Flow Prediction ◽  
Feature Based

A feature-based jet model has been proposed for use in 3D CFD prediction of turbine blade film cooling. The goal of the model is to be able to perform computationally efficient flow prediction and optimization of film-cooled turbine blades. The model reproduces in the near hole region the macro flow features of a coolant jet within a Reynolds-Averaged Navier Stokes (RANS) framework. Numerical predictions of the 3D flow through a linear transonic film-cooled turbine cascade are carried out with the model, with a low computational overhead. Different cooling holes arrangement are computed and the prediction accuracy is evaluated versus experimental data. It shown that the present model provides a reasonably good prediction of the adiabatic film-cooling effectiveness and Nusselt number around the blade. A numerical analysis of the interaction of coolant jets issuing from different rows of holes on the blade pressure side is carried out. It is shown that the upward radial migration of the flow due to the passage secondary flow structure has an impact on the spreading of the coolant and the film cooling effectiveness on the blade surface. Based on this result, a new arrangement of the cooling holes for the present case is proposed that leads to a better spanwise covering of the coolant on the blade pressure side surface.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

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.

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.

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