Full-Scale Turbine Vane Endwall Film-Cooling Effectiveness Distribution Using Pressure-Sensitive Paint Technique

2016 ◽  
Vol 138 (5) ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Andrew F Chen ◽  
Je-Chin Han ◽  
Salam Azad ◽  
Ching-Pang Lee
Keyword(s):  
Film Cooling ◽  
Full Scale ◽  
Pressure Sensitive Paint ◽  
Complex Flow ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Turbine Vane ◽  
Tunnel Design

Researchers in gas turbine field take great interest in the cooling performance on the first-stage vane because of the complex flow characteristics and intensive heat load that comes from the exit of the combustion chamber. A better understanding is needed on how the coolant flow interacts with the mainstream and the resulting cooling effect in the real engine especially for the first-stage vane. An authentic flow channel and condition should be achieved. In this study, three full-scale turbine vanes are used to construct an annular-sector cascade. The film-cooling design is attained through numerous layback fan-shaped and cylindrical holes dispersed on the vane and both endwalls. With the three-dimensional vane geometry and corresponding wind tunnel design, the true flow field can thus be simulated as in the engine. This study targets the film-cooling effectiveness on the inner endwall (hub) of turbine vane. Tests are performed under the mainstream Reynolds number 350,000; the related inlet Mach number is 0.09; and the freestream turbulence intensity is 8%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2%, 3%, and 4%) and density ratios (DR = 1.0 and 1.5), are examined. Pressure-sensitive paint (PSP) technique is utilized to capture the detail contour of film-cooling effectiveness on the inner endwall and demonstrate the coolant trace. The presented results serve as a comparison basis for other sets of vanes with different cooling designs. The results are expected to strengthen the promise of PSP technique on evaluating the film-cooling performance of the engine geometries.

Full-Scale Turbine Vane End-Wall Film-Cooling Effectiveness Distribution Using PSP Technique

2015 ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Andrew F. Chen ◽  
Je-Chin Han ◽  
Salam Azad ◽  
Ching-Pang Lee
Keyword(s):  
Film Cooling ◽  
Full Scale ◽  
Complex Flow ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Free Stream Turbulence ◽  
Turbine Vane ◽  
Tunnel Design ◽  
End Wall

Researchers in gas turbine field take great interest in the cooling performance on the first-stage vane because the complex flow characteristics and intensive heat load that comes from the exit of the combustion chamber. A better understanding is needed on how the coolant flow interacts with the mainstream and the resulting cooling effect in the real engine especially for the first-stage vane. An authentic flow channel and condition should be achieved. In this study, three full-scale turbine vanes are used to construct an annular-sector cascade. The film-cooling design is attained through numerous layback fan-shaped and cylindrical holes dispersed on the vane and both end-walls. With the three-dimensional vane geometry and corresponding wind tunnel design, the true flow field can thus be simulated as in the engine. This study targets the film-cooling effectiveness on the inner end-wall (hub) of turbine vane. Tests are performed under the mainstream Reynolds number 3.5 × 105; the related inlet Mach number is 0.09 and the free stream turbulence intensity is 8%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2%, 3%, 4%) and density ratios (DR = 1.0, 1.5) are examined. Pressure-sensitive paint (PSP) technique is utilized to capture the detail contour of film-cooling effectiveness on the inner end-wall and demonstrate the coolant trace. The presented results serve a comparison basis for other sets of vanes with different cooling designs. The results are expected to strengthen the promise of PSP technique on evaluating the film-cooling performance of the engine geometries.

Film Cooling Effectiveness Comparison on Full-Scale Turbine Vane Endwalls Using Pressure-Sensitive Paint Technique

2017 ◽  
Vol 140 (2) ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Andrew F Chen ◽  
Je-Chin Han ◽  
Salam Azad ◽  
Ching-Pang Lee
Keyword(s):  
Film Cooling ◽  
Full Scale ◽  
Manufacturing Cost ◽  
Pressure Sensitive Paint ◽  
Local Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Area Of Interest ◽  
Pressure Sensitive ◽  
Turbine Vane

Turbine vane endwalls are highly susceptible to intensive heat load due to their large exposed area and complex flow field especially for the first stage of the vane. Therefore, a suitable film cooling design that properly distributes the given amount of coolant is critical to keep the vane endwall from failure at the same time to maintain a good balance between manufacturing cost, performance, and durability. This work is focused on film cooling effectiveness evaluation on full-scale heavy-duty turbine vane endwall and the performance comparison with different film cooling pattern designs in the literature. The area of interest (AOI) of this study is on the inner endwall (hub) of turbine vane. Tests were performed in a three-vane annular sector cascade under the mainstream Reynolds number 350,000; the related inlet Mach number is 0.09 and the freestream turbulence intensity is 12%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2–4%) and density ratios (DR = 1.0, 1.5), are investigated. The conduction-error free pressure-sensitive paint (PSP) technique is utilized to evaluate the local flow behavior as well as the film cooling performance. The presented results are expected to provide the gas turbine engine designer a direct comparison between two film-hole configurations on a full-scale vane endwall under the same amount of coolant usage.

Film Cooling Effectiveness Comparison on Full-Scale Turbine Vane Endwalls Using PSP Technique

2017 ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Andrew F. Chen ◽  
Je-Chin Han ◽  
Salam Azad ◽  
Ching-Pang Lee
Keyword(s):  
Film Cooling ◽  
Flow Behavior ◽  
Full Scale ◽  
Manufacturing Cost ◽  
Complex Flow ◽  
Local Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Area Of Interest ◽  
Turbine Vane

Turbine vane endwalls are highly susceptible to intensive heat load due to their large exposed area and complex flow field especially for the first-stage of the vane. Therefore, a suitable film cooling design that properly distributes the given amount of coolant is critical to keep the vane endwall from failure at the same time to maintain a good balance between manufacturing cost, performance, and durability. This work is focused on film cooling effectiveness evaluation on full-scale heavy duty turbine vane endwall and the performance comparison with different film cooling pattern designs in the literature. The area of interest of this study is on the inner endwall (hub) of turbine vane. Tests were performed in a three-vane annular sector cascade under the mainstream Reynolds number 350,000; the related inlet Mach number is 0.09 and the freestream turbulence intensity is 12%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2%, 3%, 4%) and density ratios (DR = 1.0, 1.5) are investigated. The conduction-error free Pressure sensitive paint (PSP) technique is utilized to evaluate the local flow behavior as well as the film cooling performance. The presented results are expected to provide the gas turbine engine designer a direct comparison between two film-hole configurations on a full-scale vane endwall under the same amount of coolant usage.

Transonic Turbine Vane Endwall Film Cooling Using the Pressure-Sensitive Paint Measurement Technique

2020 ◽  
Vol 142 (8) ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Izzet Sahin ◽  
Izhar Ullah ◽  
Je-Chin Han ◽  
Alexander V. Mirzamoghadam ◽  
...  
Keyword(s):  
Shock Wave ◽  
Film Cooling ◽  
Density Ratio ◽  
Pressure Sensitive Paint ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Turbine Vane ◽  
Endwall Film Cooling

Abstract This work focuses on the parametric study of film cooling effectiveness on the turbine vane endwall under various flow conditions. The experiments were performed in a five-vane annular sector cascade facility in a blowdown wind tunnel. The controlled exit isentropic Mach numbers were 0.7, 0.9, and 1.0, from high subsonic to transonic conditions. The freestream turbulence intensity is estimated to be 12%. Three coolant-to-mainstream mass flow ratios (MFR) in the range 0.75%, 1.0%, and 1.25% are studied. N2, CO2, and Argon/SF6 mixture were used to investigate the effects of density ratio (DR), ranging from 1.0, 1.5, to 2.0. There are eight cylindrical holes on the endwall inside the passage. The pressure-sensitive paint (PSP) technique was used to capture the endwall pressure distribution for shock wave visualization and obtain the detailed film cooling effectiveness distributions. Both the high-fidelity effectiveness contour and the laterally (spanwise) averaged effectiveness were measured to quantify the parametric effect. This study will provide the gas turbine designer more insight on how the endwall film cooling effectiveness varies with different cooling flow conditions including shock wave through the endwall crossflow passage.

Film Cooling Effectiveness on a Rotating Turbine Platform Using Pressure Sensitive Paint Technique

2007 ◽  
Author(s):  
A. Suryanarayanan ◽  
B. Ozturk ◽  
M. T. Schobeiri ◽  
J. C. Han
Keyword(s):  
Film Cooling ◽  
Gas Injection ◽  
Turbine Rotor ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Coolant Injection ◽  
Purge Flow ◽  
Effects Of Rotation

Film cooling effectiveness is measured on a rotating turbine blade platform for coolant injection through discrete holes using pressure sensitive paint technique (PSP). Most of the existing literatures provide information only for stationary end-walls. The effects of rotation on the platform film cooling effectiveness are not well documented. Hence, the existing 3-stage turbine research facility at TPFL, Texas A&M University was re-designed and installed to enable coolant gas injection on the 1st stage rotor platform. Two distinct coolant supply loops were incorporated into the rotor to facilitate separate feeds for upstream cooling using stator-rotor gap purge flow and downstream discrete-hole film cooling. As a continuation of the previously published work involving stator-rotor gap purge cooling, this study investigates film cooling effectiveness on the 1st stage rotor platform due to coolant gas injection through nine discrete holes located downstream within the passage region. Film cooling effectiveness is measured for turbine rotor frequencies of 2400rpm, 2550rpm and 3000rpm corresponding to rotation numbers of Ro = 0.18, 0.19 and 0.23 respectively. For each of the turbine rotational frequencies, film cooling effectiveness is determined for average film-hole blowing ratios of Mholes = 0.5, 0.75, 1.0, 1.25, 1.5 and 2.0. To provide a complete picture of hub cooling under rotating conditions, simultaneous injection of coolant gas through upstream stator-rotor purge gap and downstream discrete film-hole is also studied. The combined tests are conducted for gap purge flow corresponding to coolant to mainstream mass flow ratio of MFR = 1% with three downstream film-hole blowing ratios of Mholes = 0.75, 1.0 and 1.25 for each of the three turbine speeds. The results for combined upstream stator-rotor gap purge flow and downstream discrete holes provide information about the optimum purge flow coolant mass, average coolant hole blowing ratios for each rotational speed and coolant injection location along the passage to obtain efficient platform film cooling.

Numerical Investigation of the Cooling Performance on the Endwall With and Without Fillet

2018 ◽  
Author(s):  
Qingzong Xu ◽  
Qiang Du ◽  
Pei Wang ◽  
Jun Liu ◽  
Guang Liu
Keyword(s):  
Secondary Flow ◽  
Turbulence Model ◽  
Film Cooling ◽  
Computational Method ◽  
Inlet Temperature ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Numerical Predictions ◽  
Turbine Vane

High inlet temperature of turbine vane increases the demand of high film cooling effectiveness. Vane endwall region was extensively cooled due to the high and flat exit temperature distribution of combustor. Leakage flow from the combustor-turbine gap was used to cool the endwall region except for preventing hot gas ingestion. Numerical predictions were conducted to investigate the flow structure and adiabatic film cooling effectiveness of endwall region in a linear cascade with vane-endwall junction fillet. The simulations were completed by solving the three-dimensional Reynolds-Averaged Navier-Stokes(RANS) equations with shear stress transport(SST) k-ω turbulence model, meanwhile, the computational method and turbulence model were validated by comparing computational result with the experiment. Three types of linear fillet with the length-to-height ratio of 0.5, 1 and 2, named fillet A, fillet B and fillet C respectively, were studied. In addition, circular fillet with radius of 2mm was compared with linear fillet B. The interrupted slot, produced by changing the way of junction of combustor and turbine vane endwall, is introduced at X/Cax = −0.2 upstream of the vane leading edge. Results showed that fillet can significantly affect the cooling performance on the endwall due to suppressing the strength of the secondary flow. Fillet C presented the best cooling performance comparing to fillet A and fillet B because a portion of the coolant which climbs to the fillet was barely affected by secondary flow. Results also showed the effect of fillet on the total pressure loss. The result indicated that only fillet A slightly decreases endwall loss.

Influence of Coolant Density on Turbine Blade Film Cooling at Transonic Cascade Flow Conditions Using the Pressure Sensitive Paint Technique

2021 ◽  
Author(s):  
Izhar Ullah ◽  
Sulaiman M. Alsaleem ◽  
Lesley M. Wright ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Pressure Side ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Blade Tip ◽  
Film Cooling Holes

Abstract This work is an experimental study of film cooling effectiveness on a blade tip in a stationary, linear cascade. The cascade is mounted in a blowdown facility with controlled inlet and exit Mach numbers of 0.29 and 0.75, respectively. The free stream turbulence intensity is measured to be 13.5 % upstream of the blade’s leading edge. A flat tip design is studied, having a tip gap of 1.6%. The blade tip is designed to have 15 shaped film cooling holes along the near-tip pressure side (PS) surface. Fifteen vertical film cooling holes are placed on the tip near the pressure side. The cooling holes are divided into a 2-zone plenum to locally maintain the desired blowing ratios based on the external pressure field. Two coolant injection scenarios are considered by injecting coolant through the tip holes only and both tip and PS surface holes together. The blowing ratio (M) and density ratio (DR) effects are studied by testing at blowing ratios of 0.5, 1.0, and 1.5 and three density ratios of 1.0, 1.5, and 2.0. Three different foreign gases are used to create density ratio effect. Over-tip flow leakage is also studied by measuring the static pressure distributions on the blade tip using the pressure sensitive paint (PSP) measurement technique. In addition, detailed film cooling effectiveness is acquired to quantify the parametric effect of blowing ratio and density ratio on a plane tip design. Increasing the blowing ratio and density ratio resulted in increased film cooling effectiveness at all injection scenarios. Injecting coolant on the PS and the tip surface also resulted in reduced leakage over the tip. The conclusions from this study will provide the gas turbine designer with additional insight on controlling different parameters and strategically placing the holes during the design process.

scholarly journals Flat Plate and Turbine Vane Film-Cooling Performance with Laid-Back Fan-Shaped Holes

2019 ◽  
Vol 4 (2) ◽  
pp. 14 ◽  
Author(s):  
Tommaso Bacci ◽  
Alessio Picchi ◽  
Bruno Facchini
Keyword(s):  
Flat Plate ◽  
Film Cooling ◽  
Density Ratio ◽  
Pressure Sensitive Paint ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Turbine Vane ◽  
Experimental Works ◽  
Lift Off

Shaped holes are considered as an effective solution to enhance gas turbine film-cooling performance, as they allow to increase the coolant mass-flux, while limiting the detrimental lift-off phenomena. A great amount of work has been carried out in past years on basic flat plate configurations while a reduced number of experimental works deals with a quantitative assessment of the influence of curvature and vane pressure gradient. In the present work PSP (Pressure Sensitive Paint) technique is used to detail the adiabatic effectiveness generated by axial shaped holes with high value of Area Ratio close to 7, in three different configurations with the same 1:1 scale: first of all, a flat plate configuration is examined; after that, the film-cooled pressure and suction sides of a turbine vane model are investigated. Tests were performed varying the blowing ratio and imposing a density ratio of 2.5 . The experimental results are finally compared to the predictions of two different correlations, developed for flat plate configurations.

Film Cooling Effectiveness on the Leading Edge Region of a Rotating Turbine Blade With Two Rows of Film Cooling Holes Using Pressure Sensitive Paint

2006 ◽  
Vol 128 (9) ◽  
pp. 879-888 ◽  
Author(s):  
Jaeyong Ahn ◽  
M. T. Schobeiri ◽  
Je-Chin Han ◽  
Hee-Koo Moon
Keyword(s):  
Rotational Speed ◽  
Film Cooling ◽  
Rotor Blade ◽  
Incidence Angle ◽  
Leading Edge ◽  
Edge Region ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive

Detailed film cooling effectiveness distributions are measured on the leading edge of a rotating gas turbine blade with two rows (pressure-side row and suction-side row from the stagnation line) of holes aligned to the radial axis using the pressure sensitive paint (PSP) technique. Film cooling effectiveness distributions are obtained by comparing the difference of the measured oxygen concentration distributions with air and nitrogen as film cooling gas respectively and by applying the mass transfer analogy. Measurements are conducted on the first-stage rotor blade of a three-stage axial turbine at 2400rpm (positive off-design), 2550rpm (design), and 3000rpm (negative off-design), respectively. The effect of three blowing ratios is also studied. The blade Reynolds number based on the axial chord length and the exit velocity is 200,000 and the total to exit pressure ratio was 1.12 for the first-stage rotor blade. The corresponding rotor blade inlet and outlet Mach numbers are 0.1 and 0.3, respectively. The film cooling effectiveness distributions are presented along with discussions on the influence of rotational speed (off design incidence angle), blowing ratio, and upstream nozzle wakes around the leading edge region. Results show that rotation has a significant impact on the leading edge film cooling distributions with the average film cooling effectiveness in the leading edge region decreasing with an increase in the rotational speed (negative incidence angle).

Film-Cooling Effectiveness on a Gas Turbine Blade Tip Using Pressure-Sensitive Paint

2005 ◽  
Vol 127 (5) ◽  
pp. 521-530 ◽  
Author(s):  
Jaeyong Ahn ◽  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Oxygen Concentration ◽  
Film Cooling ◽  
Pressure Sensitive Paint ◽  
Nitrogen Gas ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Blade Tip ◽  
Film Cooling Holes

Effects of the presence of squealer, the locations of the film-cooling holes, and the tip-gap clearance on the film-cooling effectiveness were studied and compared to those for a plane (flat) tip. The film-cooling effectiveness distributions were measured on the blade tip using the pressure-sensitive paint technique. Air and nitrogen gas were used as the film-cooling gases, and the oxygen concentration distribution for each case was measured. The film-cooling effectiveness information was obtained from the difference of the oxygen concentration between air and nitrogen gas cases by applying the mass transfer analogy. Plane tip and squealer tip blades were used while the film-cooling holes were located (a) along the camber line on the tip or (b) along the tip of the pressure side. The average blowing ratio of the cooling gas was 0.5, 1.0, and 2.0. Tests were conducted with a stationary, five-bladed linear cascade in a blow-down facility. The free-stream Reynolds number, based on the axial chord length and the exit velocity, was 1,138,000, and the inlet and the exit Mach numbers were 0.25 and 0.6, respectively. Turbulence intensity level at the cascade inlet was 9.7%. All measurements were made at three different tip-gap clearances of 1%, 1.5%, and 2.5% of blade span. Results show that the locations of the film-cooling holes and the presence of squealer have significant effects on surface static pressure and film-cooling effectiveness, with film-cooling effectiveness increasing with increasing blowing ratio.

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