SHAPE OPTIMIZATION TO IMPROVE FILM-COOLING PERFORMANCE FOR LAIDBACK FAN-SHAPED FILM COOLING HOLE ON THE SUCTION SURFACE OF A TURBINE VANE

2018 ◽  
Author(s):  
Ying Huang ◽  
Jingzhou Zhang ◽  
Chun-hua Wang ◽  
Xing-dan Zhu
Keyword(s):  
Shape Optimization ◽  
Film Cooling ◽  
Suction Surface ◽  
Cooling Performance ◽  
Turbine Vane ◽  
Cooling Hole

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.

Aero-Thermal Shape Optimization for the Discrete Film Cooling of a Turbine Airfoil

2015 ◽  
Author(s):  
Carole El Ayoubi ◽  
Wahid Ghaly ◽  
Ibrahim Hassan
Keyword(s):  
Shape Optimization ◽  
Film Cooling ◽  
Pareto Front ◽  
Multiple Objective ◽  
Suction Surface ◽  
Nsga Ii ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Rans Simulations

A multiple-objective shape optimization is implemented for two staggered rows of discrete film cooling holes on the suction surface of a turbine vane. The optimization aims to maximize the film cooling performance while minimizing the corresponding aerodynamic penalty. The cooling performance is assessed using the adiabatic film cooling effectiveness, while the aerodynamic penalty is measured with a mass-averaged total pressure loss coefficient. The conical expansion angle, the compound angle, and the length to diameter ratio of the non-expanded portion of the hole are selected as geometric design variables. The effect of varying the geometric variables on the adiabatic film cooling effectiveness and the aerodynamic penalty is analyzed using the optimization method and three-dimensional Reynolds-averaged Navier-Stokes (RANS) simulations. A non-dominated sorting genetic algorithm (NSGA-II) is coupled with an artificial neural network (ANN) to perform the multiple-objective optimization. RANS simulations are employed to construct the ANN network which produces low-fidelity predictions of the objective functions during the optimization. The Pareto front of optimum solutions is generated. Two optimum designs, denoted as the aerodynamic, and thermal optimums are chosen from the Pareto front and evaluated through RANS simulations. The optimum designs present improved performance in comparison to the reference design, which consists of cylindrical holes.

Multi-objective optimization of laidback fan-shaped film cooling hole on Turbine Vane Suction Surface

2018 ◽  
Vol 55 (4) ◽  
pp. 1181-1194 ◽  
Author(s):  
Ying Huang ◽  
Jing-Zhou Zhang ◽  
Chun-Hua Wang ◽  
Xing-Dan Zhu
Keyword(s):  
Film Cooling ◽  
Multi Objective Optimization ◽  
Suction Surface ◽  
Multi Objective ◽  
Turbine Vane ◽  
Cooling Hole

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.

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.

Computational Analysis of Conjugate Heat Transfer and Particulate Deposition on a High Pressure Turbine Vane

2009 ◽  
Author(s):  
Weiguo Ai ◽  
Thomas H. Fletcher
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Particle Method ◽  
Experimental Results ◽  
Turbine Vane ◽  
Particulate Deposition ◽  
Cooling Hole ◽  
Lagrangian Particle Method ◽  
Film Cooling Holes

Numerical computations were conducted to simulate flyash deposition experiments on gas turbine disk samples with internal impingement and film cooling using a CFD code (FLUENT). The standard k-ω turbulence model and RANS were employed to compute the flow field and heat transfer. The boundary conditions were specified to be in agreement with the conditions measured in experiments performed in the BYU Turbine Accelerated Deposition Facility (TADF). A Lagrangian particle method was utilized to predict the ash particulate deposition. User-defined subroutines were linked with FLUENT to build the deposition model. The model includes particle sticking/rebounding and particle detachment, which are applied to the interaction of particles with the impinged wall surface to describe the particle behavior. Conjugate heat transfer calculations were performed to determine the temperature distribution and heat transfer coefficient in the region close to the film-cooling hole and in the regions further downstream of a row of film-cooling holes. Computational and experimental results were compared to understand the effect of film hole spacing, hole size and TBC on surface heat transfer. Calculated capture efficiencies compare well with experimental results.

Rib Turbulator Effects on Crossflow-Fed Shaped Film Cooling Holes

2019 ◽  
Vol 141 (3) ◽  
Author(s):  
Dale W. Fox ◽  
Fraser B. Jones ◽  
John W. McClintic ◽  
David G. Bogard ◽  
Thomas E. Dyson ◽  
...  
Keyword(s):  
Film Cooling ◽  
Velocity Ratio ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Smooth Channel ◽  
Cooling Hole ◽  
Film Cooling Holes ◽  
Inlet Position ◽  
The Impact

Most studies of turbine airfoil film cooling in laboratory test facilities have used relatively large plenums to feed flow into the coolant holes. However, a more realistic inlet condition for the film cooling holes is a relatively small channel. Previous studies have shown that the film cooling performance is significantly degraded when fed by perpendicular internal crossflow in a smooth channel. In this study, angled rib turbulators were installed in two geometric configurations inside the internal crossflow channel, at 45 deg and 135 deg, to assess the impact on film cooling effectiveness. Film cooling hole inlets were positioned in both prerib and postrib locations to test the effect of hole inlet position on film cooling performance. A test was performed independently varying channel velocity ratio and jet to mainstream velocity ratio. These results were compared to the film cooling performance of previously measured shaped holes fed by a smooth internal channel. The film cooling hole discharge coefficients and channel friction factors were also measured for both rib configurations with varying channel and inlet velocity ratios. Spatially averaged film cooling effectiveness is largely similar to the holes fed by the smooth internal crossflow channel, but hole-to-hole variation due to inlet position was observed.

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