Optimization of Film Cooling Holes on the Suction Surface of a High Pressure Turbine Blade

2012 ◽  
Author(s):  
Carole El Ayoubi ◽  
Wahid Ghaly ◽  
Ibrahim Hassan
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Three Dimensional ◽  
Suction Surface ◽  
Cfd Simulations ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Flow Parameters ◽  
Film Cooling Holes

This paper aims to optimize film coolant flow parameters on the suction surface of a high-pressure gas turbine blade in order to obtain an optimum compromise between a high film cooling effectiveness and a low aerodynamic loss. An optimization algorithm coupled with three-dimensional Reynolds-averaged Navier Stokes (RANS) analysis is used to determine the optimum film cooling configuration. The VKI blade with two staggered rows of axially oriented, conically flared, film cooling holes on its suction surface is considered. Two design variables are selected; the coolant to mainstream temperature ratio and total pressure ratio. The effect of varying these coolant flow parameters on the film cooling effectiveness and the aerodynamic loss is analyzed using an optimization method and three dimensional steady CFD simulations. The optimization process involves a genetic algorithm and a response surface approximation of the artificial neural network type to provide low-fidelity predictions of the objective function. The CFD simulations are performed using the commercial software CFX. The numerical predictions of the aerodynamics and wall heat transfer are validated against experimental data. The optimization objective consists of maximizing the spatially averaged film cooling effectiveness and minimizing the aerodynamic penalty produced by film cooling. The results of this optimization are reported in terms of the aerodynamic loss and adiabatic cooling effectiveness.

Aerothermal Effect of Cavity Welding Beads on a Transonic Squealer Tip

2020 ◽  
Author(s):  
Joao Vieira ◽  
John Coull ◽  
Peter Ireland ◽  
Eduardo Romero
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Reynolds Numbers ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Blade Tip ◽  
Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using PSP Technique

2010 ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Wake ◽  
Lift Off ◽  
Film Cooling Holes

The effect of an unsteady stator wake (simulated by wake rods mounted on a spoke wheel wake generator) on the modeled rotor blade is studied using the Pressure Sensitive Paint (PSP) mass transfer analogy method. Emphasis of the current study is on the mid-span region of the blade. The flow is in the low Mach number (incompressible) regime. The suction (convex) side has simple angled cylindrical film-cooling holes; the pressure (concave) side has compound angled cylindrical film cooling holes. The blade also has radial shower-head leading edge film cooling holes. Strouhal numbers studied range from 0 to 0.36; the exit Reynolds Number based on the axial chord is 530,000. Blowing ratios range from 0.5 to 2.0 on the suction side; 0.5 to 4.0 on the pressure side. Density ratios studied range from 1.0 to 2.5, to simulate actual engine conditions. The convex suction surface experiences film-cooling jet lift-off at higher blowing ratios, resulting in low effectiveness values. The film coolant is found to reattach downstream on the concave pressure surface, increasing effectiveness at higher blowing ratios. Results show deterioration in film cooling effectiveness due to increased local turbulence caused by the unsteady wake, especially on the suction side. Results also show a monotonic increase in film-cooling effectiveness on increasing the coolant to mainstream density ratio.

Interactions Between the Combustor Swirl and the High Pressure Stator of a Turbine

2012 ◽  
Author(s):  
Lucas Giller ◽  
Heinz-Peter Schiffer
Keyword(s):  
High Pressure ◽  
Flow Field ◽  
Film Cooling ◽  
Field Measurements ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Film Cooling Holes ◽  
The Impact

The interaction between the strongly swirling combustor outflow and the high pressure turbine nozzle guide vanes were investigated at the cascade test rig at Technische Universität Darmstadt. The test section of the rig consists of six swirl generators and five cascade vanes. The three middle vanes are equipped with film cooling holes at the leading edges. The swirler nozzles are aligned with the center of the cascade passages. The operating settings are defined by the swirl number, the distance between the swirler nozzles and the vanes, the blowing ratio and the radial angle of the film cooling holes. Flow field measurements using PIV downstream of the swirlers and five hole probe measurements at the inlet and outlet plane of the cascade were accomplished. Measurements using the ammonia diazo technique to determine the adiabatic film cooling effectiveness on the surface of the center cascade vane were also carried out. It is shown that a swirling inflow leads to a strong alteration of the flow field and the losses in the passages in comparison to an axial inflow. Furthermore, the impact of the swirl on the formation of the cooling film and it’s adiabatic film cooling effectiveness is presented.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using Pressure Sensitive Paint Technique

2012 ◽  
Vol 134 (8) ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Suction Side ◽  
Pressure Sensitive Paint ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Unsteady Wake ◽  
Film Cooling Holes

The effect of an unsteady stator wake (simulated by wake rods mounted on a spoke-wheel wake generator) on the modeled rotor blade is studied using the pressure sensitive paint (PSP) mass-transfer analogy method. Emphasis of the current study is on the midspan region of the blade. The flow is in the low Mach number (incompressible) regime. The suction (convex) side has simple angled cylindrical film-cooling holes; the pressure (concave) side has compound angled cylindrical film-cooling holes. The blade also has radial shower-head leading edge film-cooling holes. Strouhal numbers studied range from 0 to 0.36; the exit Reynolds number based on the axial chord is 530,000. Blowing ratios range from 0.5 to 2.0 on the suction side and 0.5 to 4.0 on the pressure side. Density ratios studied range from 1.0 to 2.5, to simulate actual engine conditions. The convex suction surface experiences film-cooling jet lift-off at higher blowing ratios, resulting in low effectiveness values. The film coolant is found to reattach downstream on the concave pressure surface, increasing effectiveness at higher blowing ratios. Results show deterioration in film-cooling effectiveness due to increased local turbulence caused by the unsteady wake, especially on the suction side. Results also show a monotonic increase in film-cooling effectiveness on increasing the coolant to mainstream density ratio.

Effect of Upstream Wake on Passage Flow and Tip Film Cooling Characteristics

2012 ◽  
Author(s):  
Jin Wang ◽  
Bengt Sundén ◽  
Min Zeng ◽  
Qiu-Wang Wang
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Flow Distribution ◽  
Tip Leakage Flow ◽  
Delta Wings ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Passage Vortex ◽  
Film Cooling Holes

Three-dimensional simulations of the squealer tip on the GE-E3 blade with eight film cooling holes were carried out. To form the wake by the trailing edges of the stator vanes, cylindrical rods and delta wings were placed upstream of the blades. The rods were placed according to three positions, and the influence on the film cooling effectiveness was calculated. Because delta wings were placed upstream of the blades to generate in the vane passage, the passage flow also was investigated. However, the passage vortex generated by the delta wings had a profound effect on the passage flow distribution. For the squealer tip, the cavity contributes to the improvement of the cooling effect in the tip zone. The passage flow and the tip leakage flow influenced by cylindrical rods and delta wings were analyzed using numerical simulations with the blowing ratio of M = 0.5. In addition, calculations with and without cylindrical rods and delta wings were performed and then comparisons were enabled. It was found that the vortex created by delta wings made the passage flow more turbulent and the result indicates a slight effect on the film cooling effectiveness in the tip gap.

Optimal Arrangement of the Film Cooling Holes Considering the Manufacturing Tolerance for High Pressure Turbine Nozzle

2016 ◽  
Author(s):  
Sanga Lee ◽  
Dong-Ho Rhee ◽  
Kwanjung Yee
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Manufacturing Tolerance ◽  
Cooling Hole ◽  
Film Cooling Holes

In spite of a myriad of researches on the optimal shape of film cooling holes, only a few attempts have been made to optimize the hole arrangement for film cooling so far. Moreover, although the general scale of film cooling hole is so small that manufacturing tolerance has substantial effects on the cooling performance of turbine, the researches on this issue are even scarcer. If it is possible to obtain optimal hole arrangement which not only improve the film cooling performance but also is robust to the manufacturing tolerance, then overall cooling performance of a turbine would become more reliable and useful from the practical point of view. To this end, the present study proposed a robust design optimization procedure which takes the manufacturing uncertainties into account. The procedure was subsequently applied to the film cooling holes on high pressure turbine nozzle pressure side to obtain the robust array shape under the uncertainty of the manufacturing tolerance. First, the array of the holes was parameterized by 5 design variables using the newly suggested shape functions, and 2 representative factors were considered for the manufacturing tolerance of the film cooling hole. Probabilistic process that consists of Kriging surrogate model and Monte Carlo Simulation with descriptive sampling method was coupled with the design optimization process using Genetic Algorithm. Through this, film cooling hole array which shows the high performance, yet robust to the manufacturing tolerance was obtained, and the effects of the manufacturing tolerance on the cooling performance was carefully investigated. As a result, the region where the film cooling effectiveness is noticeable, as well as the maximum width of the variation of the film cooling effectiveness were reduced through optimization, and it is also confirmed that the tolerance of the holes near the leading edge is more influential to the cooling performance because the film cooling effectiveness is more sensitive to the manufacturing tolerance of the leading edge than that of the trailing edge.

Tripod Hole Geometry Performance for a Vane Suction Surface Near Throat Location

2013 ◽  
Author(s):  
Sridharan Ramesh ◽  
Chris LeBlanc ◽  
Srinath V. Ekkad ◽  
Mary Anne Alvin
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Throat Region ◽  
Blowing Ratio ◽  
Hole Geometry ◽  
Film Cooling Holes ◽  
Cooling Air

Film cooling performance depends strongly on the hole exit geometry, blowing ratio, and hole location. The goal of this study is to evaluate film cooling geometries that can provide better protection over the suction surface of the airfoil beyond the throat region. This study compares the performance of standard cylindrical; fan-shaped (10° flare/laidback); tripod hole geometry (15° breakout angle); and tripod holes with shaped exits (5° flare on 15° tripod). Film cooling holes are located just upstream of the throat region on the suction side of an airfoil. The airfoil is a scaled up first stage vane from GE E3 engine and is mounted on a low speed linear cascade wind tunnel. A range of blowing ratios from 0.5 to 2.0 was covered for a cylindrical hole, while ensuring all other hole geometries run under similar mass flow rate conditions. Steady state IR (Infra-Red) technique was employed to measure adiabatic film cooling effectiveness. Results show that the tripod holes with and without shaped exits provide much higher film effectiveness than cylindrical and slightly higher effectiveness than shaped exit holes using 50% lesser cooling air while operating at the same blowing ratios. Effectiveness values up to 0.2–0.25 are seen 40-hole diameters downstream for the tripod hole configurations thus providing cooling in the important trailing edge portion of the airfoil.

AEROTHERMAL EFFECT OF CAVITY WELDING BEADS ON A TRANSONIC SQUEALER TIP

2021 ◽  
pp. 1-12
Author(s):  
Joao Vieira ◽  
John D Coull ◽  
Peter Ireland ◽  
Eduardo Romero
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Performance Measurements ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Blade Tip ◽  
Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

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.

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