Experimental Simulation of Contaminant Deposition on a Film Cooled Turbine Airfoil Leading Edge

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Jason E. Albert ◽  
David G. Bogard
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Coal Ash ◽  
Leading Edge ◽  
Experimental Simulation ◽  
Model Surface ◽  
Blowing Ratio ◽  
The Difference ◽  
Wax Deposits ◽  
Turbine Airfoils

A significant challenge of utilizing coal-derived synthetic fuels for gas turbine engines is mitigating the adverse effects of fuel-born contaminant deposits on film cooled turbine surfaces. A new experimental technique has been developed that simulates the key physical, but not the chemical, aspects of coal ash deposition on film cooled turbine airfoil leading edges in order to better understand the interaction between film cooling and deposition and to produce improved film cooling designs. In this large-scale wind tunnel facility, the depositing contaminants were modeled with atomized molten wax droplets sized to match the Stokes numbers of coal ash particles in the engine conditions. The sticking mechanism of the molten contaminants to the turbine surfaces was modeled by ensuring the wax droplets remained somewhat molten when they arrived at the cooled model surface. The airfoil model and wax deposits had thermal conductivities such that they matched the Biot numbers of clean and fouled turbine airfoils at engine conditions. The behavior of the deposit growth was controlled by adjusting the mainstream, coolant, and wax solidification temperatures. Simulated deposits were created for a range of test durations, film cooling blowing ratios, and controlling temperatures. Inspection of the resulting deposits revealed aspects of the flow field that augment and suppress deposition. Deposit thickness was found to increase in time until an equilibrium thickness was attained. Blowing ratio and the difference between mainstream and wax solidification temperatures strongly affected characteristics of the deposits. Model surface temperatures greatly reduced under the deposits as they developed.

Experimental Simulation of Contaminant Deposition on a Film Cooled Turbine Airfoil Leading Edge

2009 ◽  
Author(s):  
Jason E. Albert ◽  
Kelly J. Keefe ◽  
David G. Bogard
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Coal Ash ◽  
Leading Edge ◽  
Experimental Simulation ◽  
Model Surface ◽  
Blowing Ratio ◽  
The Difference ◽  
Wax Deposits ◽  
Turbine Airfoils

A significant challenge of utilizing coal-derived synthetic fuels for gas turbine engines is mitigating the adverse effects of fuel-born contaminant depositions on film cooled turbine surfaces. A new experimental technique has been developed that simulates the key physical, but not the chemical, aspects of coal ash deposition on film cooled turbine airfoil leading edges in order to better understand the interaction between film cooling and deposition and to produce improved film cooling designs. In this large-scale wind tunnel facility, the depositing contaminants were simulated by atomized molten wax droplets sized to match the Stokes numbers of coal ash particles in the engine conditions. The sticking mechanism of the molten contaminants to the turbine surfaces was simulated by ensuring the wax droplets remained somewhat molten when they arrived at the cooled model surface. The model and wax deposits had thermal conductivities such that they matched the Biot numbers of clean and fouled turbine airfoils at engine conditions. The behavior of the deposition growth was controlled by adjusting the mainstream, coolant, and wax solidification temperatures. Simulated depositions were created for a range of test durations, film cooling blowing ratios, and controlling temperatures. Inspection of the resulting depositions revealed aspects of the flow field that augment and suppress deposition. Deposition thickness was found to increase in time until a quasi-steady thickness was attained. Blowing ratio and the difference between mainstream and wax solidification temperatures strongly affected characteristics of the depositions. Model surface temperatures greatly reduced under the depositions as they developed.

Effects of Surface Deposition, Hole Blockage, and TBC Spallation on Vane Endwall Film-Cooling

2006 ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Large Scale ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Blowing Ratio ◽  
Barrier Coating ◽  
Cooling Hole ◽  
Endwall Film Cooling ◽  
Alternate Fuels

With the increase in usage of gas turbines for power generation and given that natural gas resources continue to be depleted, it has become increasingly important to search for alternate fuels. One source of alternate fuels is coal derived synthetic fuels. Coal derived fuels, however, contain traces of ash and other contaminants that can deposit on vane and turbine surfaces affecting their heat transfer through reduced film-cooling. The endwall of a first stage vane is one such region that can be susceptible to depositions from these contaminants. This study uses a large-scale turbine vane cascade in which the following effects on film-cooling adiabatic effectiveness were investigated in the endwall region: the effect of near-hole deposition, the effect of partial film-cooling hole blockage, and the effect of spallation of a thermal barrier coating. The results indicated that deposits near the hole exit can sometimes improve the cooling effectiveness at the leading edge, but with increased deposition heights the cooling deteriorates. Partial hole blockage studies revealed that the cooling effectiveness deteriorates with increases in the number of blocked holes. Spallation studies showed that for a spalled endwall surface downstream of the leading edge cooling row, cooling effectiveness worsened with an increase in blowing ratio.

The Experimental and Numerical Research for Linear Cascade Film Cooling With Different Hole Shapes

2010 ◽  
Author(s):  
Yi Lu ◽  
Yinyi Hong ◽  
Zhirong Lin ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Linear Cascade ◽  
Blowing Ratio ◽  
Cylindrical Holes

Detailed film cooling effectiveness distributions were experimentally obtained on a turbine vane platform within a linear cascade. Testing was done in a large scale five-vane cascade with low freestream Renolds number condition 634,000 based on the axial chord length and the exit velocity. The detailed film-cooling effectiveness distributions on the platform were obtained using pressure sensitive paint technique. Two film-cooling hole configurations, cylindrical and fan-shaped, were used to cool the vane surface with two rows on pressure side, two rows on suction side and three rows on leading edge. For cylindrical holes, the blowing ratio of the coolant through the discrete cooling holes on pressure side and suction side ranged from 0.3 to 1.5 (based on the inlet mainstream velocity) while the blowing ratio ranging from 0.15 to 1.5 on leading edge; for fan-shaped holes, the four blowing ratios were 0.5, 1.0, 1.5 and 2.0. Results showed that average film-cooling effectiveness decreased with increasing blowing rate for the cylindrical holes, while the fan-shaped passage showed increased film-cooling effectiveness with increasing blowing ratio, indicating the fan-shaped cooling holes helped to improve film-cooling effectiveness by reducing overall jet liftoff. Fan-shaped holes improved average film-cooling effectiveness by 93.2%, 287.6% and 489.6% on pressure side, −4.1%, 27.9% and 78.2% on suction side over cylindrical holes at the blowing ratio of 0.5, 1.0 and 1.5 respectively. Numerical results were used to analyze the details of the flow and heat transfer on the cooling area with two turbulence models. Results demonstrated that tendency of the film cooling effectiveness distribution of numerical calculation and experimental measurement was generally consistent at different blowing ratio.

Experimental Simulation of Turbine Airfoil Leading Edge Film Cooling

1988 ◽  
Vol 110 (2) ◽  
pp. 226-232 ◽  
Author(s):  
A. R. Wadia ◽  
D. A. Nealy
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Experimental Simulation ◽  
Coolant Temperature ◽  
Film Cooling Effectiveness ◽  
Surface Heat Transfer Coefficient ◽  
Free Stream Mach Number ◽  
Turbine Airfoils

Leading edge showerhead cooling designs represent an important feature of certain classes of high-temperature turbine airfoils. This paper outlines a methodology for predicting the surface temperatures of showerhead designs with spanwise injection through an array of discrete holes. The paper describes a series of experiments and analyses on scaled cylinder models with injection through holes inclined at 20, 30, 45, and 90 deg for typical radial and circumferential spacing-to-diameter ratios of 10 and 4, respectively. The experiments were conducted in a wind tunnel on several stainless steel test specimens in which flow and heat transfer parameters were measured over the simulated airfoil leading edge surfaces. Based on the experiments, an engineering design model is proposed that treats the gas-to-surface heat transfer coefficient with film cooling in a manner suggested by a recent Purdue–NASA investigation and includes the important contribution of upstream (coolant inlet face) heat transfer. The experiments suggest that the averaged film cooling effectiveness in the showerhead region is primarily influenced by the inclination of the injection holes. The effectiveness parameter is not strongly affected by variations in coolant-to-gas stream pressure ratio, free-stream Mach number, gas-to-coolant temperature ratio, and gas stream Reynolds number. The model is employed to determine (inferentially) the average Stanton number reduction parameter for a series of pressure ratios varying from 1.004 to 1.3, Mach numbers ranging from 0.1 to 0.2, temperature ratios between 1.6 and 2.0, and Reynolds numbers ranging from 3.5×104 to 9.0×104.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

2017 ◽  
Author(s):  
Mingjie Zhang ◽  
Nian Wang ◽  
Andrew F. Chen ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Profile ◽  
Elliptical Cylinders ◽  
Blade Leading Edge

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio and blowing ratio are studied. Computational simulations are performed using the realizable k-ε turbulence model. Effectiveness obtained by CFD simulations are compared with experiments. Three leading edge profiles, including one semi-cylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semi-cylinder model, shaped holes are located at 0 degrees (stagnation line) and ± 30 degrees. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,900 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile and on turbine blade leading edge region film cooling with shaped-hole designs.

Numerical Study on the Influence of Trench Width on Film Cooling Characteristics of Double-Wave Trench

2017 ◽  
Author(s):  
Bo-lun Zhang ◽  
Li Zhang ◽  
Hui-ren Zhu ◽  
Jian-sheng Wei ◽  
Zhong-yi Fu
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Trench Width

Film cooling performance of the double-wave trench was numerically studied to improve the film cooling characteristics. Double-wave trench was formed by changing the leading edge and trailing edge of transverse trench into cosine wave. The film cooling characteristics of transverse trench and double-wave trench were numerically studied using Reynolds Averaged Navier Stokes (RANS) simulations with realizable k-ε turbulence model and enhanced wall treatment. The film cooling effectiveness and heat transfer coefficient of double-wave trench at different trench width (W = 0.8D, 1.4D, 2.1D) conditions are investigated, and the distribution of temperature field and flow field were analyzed. The results show that double-wave trench effectively improves the film cooling effectiveness and the uniformity of jet at the downstream wall of the trench. The span-wise averaged film cooling effectiveness of the double-wave trench model increases 20–63% comparing with that of the transverse trench at high blowing ratio. The anti-counter-rotating vortices which can press the film on near-wall are formed at the downstream wall of the double-wave trench. With the double-wave trench width decreasing, the film cooling effectiveness gradually reduces at the hole center-line region of the downstream trench. With the increase of the blowing ratio, the span-wise averaged heat transfer coefficient increases. The span-wise averaged heat transfer coefficient of the double-wave trench with 0.8D and 2.1D trench width is higher than that of the double-wave trench with 1.4D trench width at the high blowing ratio conditions.

Experimental Study of the Effects of an Oscillating Approach Flow on Overall Cooling Performance of a Simulated Turbine Blade Leading Edge

2009 ◽  
Author(s):  
Ross Johnson ◽  
Jonathan Maikell ◽  
David Bogard ◽  
Justin Piggush ◽  
Atul Kohli ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Stagnation Line ◽  
Cooling Performance ◽  
Blowing Ratio ◽  
Oscillation Frequencies ◽  
Overall Effectiveness ◽  
Blade Leading Edge

When a turbine blade passes through wakes from upstream vanes it is subjected to an oscillation of the direction of the approach flow resulting in the oscillation of the position of the stagnation line on the leading edge of the blade. In this study an experimental facility was developed that induced a similar oscillation of the stagnation line position on a simulated turbine blade leading edge. The overall effectiveness was evaluated at various blowing ratios and stagnation line oscillation frequencies. The location of the stagnation line on the leading edge was oscillated to simulate a change in angle of attack between α = ± 5° at a range of frequencies from 2 to 20 Hz. These frequencies were chosen based on matching a range of Strouhal numbers typically seen in an engine due to oscillations caused by passing wakes. The blowing ratio was varied between M = 1, M = 2, and M = 3. These experiments were carried out at a density ratio of DR = 1.5 and mainstream turbulence levels of Tu ≈ 6%. The leading edge model was made of high conductivity epoxy in order to match the Biot number of an actual engine airfoil. Results of these tests showed that the film cooling performance with an oscillating stagnation line was degraded by as much as 25% compared to the performance of a steady flow with the stagnation line aligned with the row of holes at the leading edge.

Leading Edge Film Cooling: Computations and Experiments Including Density Effects

1996 ◽  
Author(s):  
Pingfan He ◽  
Dragos Licu ◽  
Martha Salcudean ◽  
Ian S. Gartshore
Keyword(s):  
Experimental Data ◽  
Film Cooling ◽  
Stokes Equations ◽  
Leading Edge ◽  
Variable Density ◽  
Navier Stokes ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

The effect of varying coolant density on film cooling effectiveness for a turbine blade-model was numerically investigated and compared with experimental data. This model had a semi-circular leading edge with four rows of laterally-inclined film cooling orifices positioned symmetrically about the stagnation line. A curvilinear coordinate-based CFD code was developed and used for the numerical investigation. The code used a domain segmentation strategy in conjunction with general curvilinear grids to model the complex blade configuration. A multigrid method was used to accelerate the convergence rate. The time-averaged, variable-density, Navier-Stokes equations together with the energy or scalar equation were solved. Turbulence closure was attained by the standard k–ε model with a near-wall k model. Either air or CO2 was used as coolant in three cases of injection through single rows and alternatively staggered double raws of holes. Two different blowing rates were investigated in each case and compared with experimental data. The experimental results were obtained using a wind tunnel model, and the mass/heat analogy was used to determine the film cooling effectiveness. The higher density of the carbon dioxide coolant (approximately 1.5 times the density of air) in the isothermal mass injection experiments, was used to simulate the effects of injection of a colder air in the corresponding adiabatic heat transfer situation. Good agreement between calculated and measured film cooling effectiveness was found for low blowing ratio M ≤ 0.5 and the effect of density was not significant. At higher blowing ratio M > 1 the calculations consistently overpredict the measured values of film cooling effectiveness.

Numerical Analysis on the Leading Edge Film Cooling of Bifurcation Holes for Gas Turbine Blade

2021 ◽  
Author(s):  
Zhonghao Tang ◽  
Gongnan Xie ◽  
Honglin Li ◽  
Wenjing Gao ◽  
Chunlong Tan ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Turbulence Models ◽  
Leading Edge ◽  
Spanwise Direction ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Cylindrical Holes ◽  
Lift Off

Abstract Film cooling performance of the cylindrical film holes and the bifurcated film holes on the leading edge model of the turbine blade are investigated in this paper. The suitability of different turbulence models to predict local and average film cooling effectiveness is validated by comparing with available experimental results. Three rows of holes are arranged in a semi-cylindrical model to simulate the leading edge of the turbine blade. Four different film cooling structures (including a cylindrical film holes and other three different bifurcated film holes) and four different blowing ratios are studied in detail. The results show that the film jets lift off gradually in the leading edge area as the blowing ratio increases. And the trajectory of the film jets gradually deviate from the mainstream direction to the spanwise direction. The cylindrical film holes and vertical bifurcated film holes have better film cooling effectiveness at low blowing ratio while the other two transverse bifurcated film holes have better film cooling effectiveness at high blowing ratio. And the film cooling effectiveness of the transverse bifurcated film holes increase with the increasing the blowing ratio. Additionally, the advantage of transverse bifurcated holes in film cooling effectiveness is more obvious in the downstream region relative to the cylindrical holes. The Area-Average film cooling effectiveness of transverse bifurcated film holes is 38% higher than that of cylindrical holes when blowing ratio is 2.

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