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.

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.

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.

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.

Effects of Obstructions and Surface Roughness on Film Cooling Effectiveness With and Without a Transverse Trench

2008 ◽  
Vol 131 (1) ◽  
Author(s):  
Ruwan P. Somawardhana ◽  
David G. Bogard
Keyword(s):  
Surface Roughness ◽  
Film Cooling ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Surface Conditions ◽  
Turbine Vane ◽  
Cylindrical Holes ◽  
Better Than

Recent studies have shown that film cooling with holes embedded in a shallow trench significantly improves cooling performance. In this study, the performance of shallow trench configurations was investigated for simulated deteriorated surface conditions, i.e., increased surface roughness and near-hole obstructions. Experiments were conducted on the suction side of a scaled-up simulated turbine vane. Results from the study indicated that as much as 50% degradation occurred with upstream obstructions, but downstream obstructions actually enhanced film cooling effectiveness. However, the transverse trench configuration performed significantly better than the traditional cylindrical holes, both with and without obstructions and almost eliminated the effects of both surface roughness and obstructions.

Turbine Vane Endwall Film Cooling With Slashface Leakage and Discrete Hole Configuration

2017 ◽  
Vol 139 (6) ◽  
Author(s):  
Nafiz H. K. Chowdhury ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han ◽  
Luzeng Zhang ◽  
Hee-Koo Moon
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Secondary Flows ◽  
Blow Down ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Free Stream Turbulence ◽  
Leakage Flows ◽  
Turbine Vane

Turbine vanes are typically assembled as a section containing single or double airfoil units in an annular pattern. First stage guide vane assembly results in two common mating interfaces: a gap between combustor and vane endwall and another resulted from the adjacent sections, called slashface. High pressure coolant could leak through these gaps to reduce the ingestion of hot gas and achieve certain cooling benefit. As vane endwall region flow field is already very complicated due to highly three-dimensional secondary flows, then a significant influence on endwall cooling can be expected due to the gap leakage flows. To determine the effect of leakage flows from those gaps, film cooling effectiveness distributions were measured using pressure sensitive paint (PSP) technique on the endwall of a scaled up, midrange industrial turbine vane geometry with the multiple rows of discrete film cooling (DFC) holes inside the passages. Experiments were performed in a blow-down wind tunnel cascade facility at the exit Mach number of 0.5 corresponding to Reynolds number of 3.8 × 105 based on inlet conditions and axial chord length. Passive turbulence grid was used to generate free-stream turbulence (FST) level about 19% with an integral length scale of 1.7 cm. Two parameters, coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR), were studied. The results are presented as two-dimensional film cooling effectiveness distribution on the vane endwall surface with the corresponding spanwise averaged values along the axial direction.

Turbine Vane Endwall Film Cooling Comparison From Five Film-Hole Design Patterns and Three Upstream Injection Angles

2019 ◽  
Vol 11 (3) ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Izzet Sahin ◽  
Nian Wang ◽  
Je-Chin Han ◽  
Hongzhou Xu ◽  
...  
Keyword(s):  
Film Cooling ◽  
Design Patterns ◽  
Density Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Operation Conditions ◽  
Turbine Vane ◽  
End Wall ◽  
Endwall Cooling

Abstract The effects of upstream injection angle on film cooling effectiveness of a turbine vane end wall with various endwall film-hole designs were examined by applying pressure-sensitive paint (PSP) measurement technique. As the leakage flow from the slot between the combustor and the turbine vane is not considered an active source to protect the vane endwall in certain engine designs, discrete cylindrical holes are implemented near the slot to create an additional controllable upstream film to cool the vane end wall. Three potential injection angles were studied: 30 deg, 40 deg, and 50 deg. To explore the optimum endwall cooling design, five different film-hole patterns were tested: axial row, cross row, cluster, midchord row, and downstream row. Experiments were conducted in a four-passage linear cascade facility in a blowdown wind tunnel at the exit isentropic Mach number of 0.5 corresponding to inlet Reynolds number of 380,000 based on turbine vane axial chord length. A freestream turbulence intensity of 19% with an integral length scale of 1.7 cm was generated at the cascade inlet plane. Detailed film cooling effectiveness for each design was analyzed and compared at the design operation conditions (coolant mass flow ratio (MFR) 1% and density ratio 1.5). The results are presented in terms of high-fidelity film effectiveness contours and laterally (spanwise) averaged effectiveness. This paper will provide the gas turbine designers valuable information on how to select the best endwall cooling pattern with minimum cooling air consumption over a range of upstream injection angle.

Effects of Obstructions and Surface Roughness on Film Cooling Effectiveness With and Without a Transverse Trench

2007 ◽  
Author(s):  
Ruwan P. Somawardhana ◽  
David G. Bogard
Keyword(s):  
Surface Roughness ◽  
Film Cooling ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Surface Conditions ◽  
Turbine Vane ◽  
Cylindrical Holes ◽  
Better Than

Recent studies have shown that film cooling with holes imbedded in a shallow trench significantly improve cooling performance. In this study, the performance of shallow trench configurations were investigated for simulated deteriorated surface conditions, i.e. increased surface roughness and near hole obstructions. Experiments were conducted on the suction side of a scaled-up simulated turbine vane. Results from the study indicated that as much as 50% degradation occurred with upstream obstructions, but downstream obstructions actually enhanced film cooling effectiveness. However, the transverse trench configuration performed significantly better than the traditional cylindrical holes, both with and without obstructions and almost eliminated the effects of both surface roughness and obstructions.

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.

Aero-Thermal Optimization and Experimental Verification for the Discrete Film Cooling of a Turbine Airfoil

2013 ◽  
Author(s):  
Carole El Ayoubi ◽  
Othman Hassan ◽  
Wahid Ghaly ◽  
Ibrahim Hassan
Keyword(s):  
Film Cooling ◽  
Total Pressure ◽  
Pareto Front ◽  
Optimal Solutions ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Flow Parameters ◽  
Turbine Vane

The optimization aims to maximize the film cooling performance while minimizing the corresponding aerodynamic penalty. The film cooling performance is assessed using the adiabatic film cooling effectiveness, while the aerodynamic penalty is measured with a mass-averaged total pressure loss coefficient. Two design variables are selected; the coolant to mainstream temperature ratio and total pressure ratio. Two staggered rows of discrete cylindrical film cooling holes on the suction surface of a turbine vane are considered. The effect of varying the coolant flow parameters on the adiabatic film cooling effectiveness and the aerodynamic loss is analyzed using the optimization method and three-dimensional Reynolds-averaged Navier-Stokes (RANS) simulations. The CFD predictions of the adiabatic film cooling effectiveness and aerodynamic performance are assessed and validated against corresponding experimental measurements. The optimal solutions are reproduced in the experimental facility and the Pareto front is substantiated with experimental data. A non-dominated sorting genetic algorithm (NSGA-II) is coupled with an artificial neural network (ANN) to perform a multiple objective optimization of the film coolant flow parameters on the suction surface of a high pressure gas turbine vane. The numerical predictions are employed to construct the artificial neural network that produces low-fidelity predictions of the objectives during the optimization. The Pareto front of optimal solutions is generated by the optimization methodology.

The Effects of Obstructions on Film Cooling Effectiveness on the Suction Side of a Gas Turbine Vane

2006 ◽  
Author(s):  
Patricia Demling ◽  
David G. Bogard
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Temperature Profiles ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness

The effects of obstructions on film cooling performance on a scaled-up 1st stage turbine vane will be discussed. Experimental results show that obstructions located upstream or inside of a film cooling hole will degrade adiabatic effectiveness up to 80% of the levels found with no obstructions. Downstream obstructions had little effect on performance. The location where the upstream obstructions ceased to degrade adiabatic effectiveness was determined and temperature profiles were constructed to determine how the upstream obstructions were affecting the mainstream and coolant flow.

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