Effect of Thermal Barrier Coating and Gas Radiation on Film Cooling of a Corrugated Surface

In this work, a numerical study is conducted to investigate film cooling of a corrugated surface. A conjugate heat transfer analysis is carried out, accounting for the presence of thermal barrier coating (TBC) and gas radiation. The Mach number of mainstream flow is maintained at Ma = 0.6, while cold stream Mach number is varied from 0.3 to 0.58, and density ratio is kept 4. From this study, it is observed that the overall film cooling effectiveness increases by a value ranging from 0.10 to 0.15 with the use of TBC. The hot side metallic wall temperature increases in the range of 100–150 °C when the effect of gas radiation is considered. It is also found that the film cooling effectiveness decreases with decrease in the cold side Mach number.

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Effects of Surface Deposition, Hole Blockage, and Thermal Barrier Coating Spallation on Vane Endwall Film Cooling

Journal of Turbomachinery ◽

10.1115/1.2720485 ◽

2006 ◽

Vol 129 (3) ◽

pp. 599-607 ◽

Cited By ~ 27

Author(s):

N. Sundaram ◽

K. A. Thole

Keyword(s):

Thermal Barrier Coating ◽

Film Cooling ◽

Gas Turbines ◽

Large Scale ◽

Leading Edge ◽

Thermal Barrier ◽

Cooling Effectiveness ◽

Barrier Coating ◽

Cooling Hole ◽

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.

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Film Cooling Measurements for Cratered Cylindrical Inclined Holes

Journal of Turbomachinery ◽

10.1115/1.2950055 ◽

2008 ◽

Vol 131 (1) ◽

Cited By ~ 21

Author(s):

Yiping Lu ◽

Alok Dhungel ◽

Srinath V. Ekkad ◽

Ronald S. Bunker

Keyword(s):

Heat Transfer ◽

Thermal Barrier Coating ◽

Film Cooling ◽

Dynamics Simulation ◽

Thermal Barrier ◽

Base Line ◽

Freestream Velocity ◽

Film Cooling Effectiveness ◽

Barrier Coating ◽

Cylindrical Holes

Film cooling performance is studied for cylindrical holes embedded in craters. Different crater geometries are considered for a typical crater depth. Cratered holes may occur when blades are coated with thermal barrier coating layers by masking the hole area during thermal barrier coating (TBC) spraying, resulting in a hole surrounded by a TBC layer. The film performance and behavior is expected to be different for the cratered holes compared to standard cylindrical holes. Detailed heat transfer coefficient and film effectiveness measurements are obtained simultaneously using a single test transient IR thermography technique. The study is performed at a single mainstream Reynolds number based on freestream velocity and film-hole diameter of 11,000 at four different coolant-to-mainstream blowing ratios of 0.5, 1.0, 1.5, and 2.0. The results show that film cooling effectiveness is slightly enhanced by cratering of holes, but a substantial increase in heat transfer enhancement negates the benefits of higher film effectiveness. Three different crater geometries are studied and compared to a base line flush cylindrical hole, a trenched hole, and a typical diffuser shaped hole. Computational fluid dynamics simulation using FLUENT was also performed to determine the jet-mainstream interactions associated with the experimental surface measurements.

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The Application of Adiabatic Film Cooling Effectiveness to Non-Adiabatic Blade Cooling Design

1983 Joint Power Generation Conference: GT Papers ◽

10.1115/83-jpgc-gt-1 ◽

1983 ◽

Author(s):

C. P. Lee ◽

J. C. Han

Keyword(s):

Heat Transfer ◽

Film Cooling ◽

Heat Transfer Analysis ◽

Adiabatic Wall ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

Transfer Parameter ◽

Adiabatic Effectiveness ◽

C 32 ◽

Heat Transfer Parameter

The effect of heat transfer on film cooling has been studied analytically. The proposed model shows that the non-adiabatic film cooling effectiveness will increase with increasing of the heat transfer parameter, Ū / (ρVCp)2, on the convex, the flat and the concave walls over the entire range of film cooling parameter, X/MS. On the convex wall with a blowing rate, M, of 0.51 and a heat transfer parameter of 10−3 at the typical engine conditions, the non-adiabatic effectiveness can be higher than the adiabatic effectiveness by 45% at a film cooling parameter of 103; while the film temperature can be lower than the adiabatic wall by 18°C (32°F) at a dimensionless distance of 500. The model can be extended and applied to the heat transfer analysis for any kind of turbine blade with film cooling.

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Adiabatic Wall Effectiveness Measurements of Film-Cooling Holes With Expanded Exits

Volume 3: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/97-gt-164 ◽

1997 ◽

Cited By ~ 25

Author(s):

M. Gritsch ◽

A. Schulz ◽

S. Wittig

Keyword(s):

Mach Number ◽

Film Cooling ◽

Thermal Protection ◽

Lateral Spreading ◽

Cylindrical Hole ◽

Adiabatic Wall ◽

Camera System ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

Cooling Hole

This paper presents detailed measurements of the film-cooling effectiveness for three single, scaled-up film-cooling hole geometries. The hole geometries investigated include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fanshaped and a laidback fanshaped hole). The flow conditions considered are the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the blowing ratio (up to 2). The coolant-to-mainflow temperature ratio is kept constant at 0.54. The measurements are performed by means of an infrared camera system which provides a two-dimensional distribution of the film-cooling effectiveness in the nearfield of the cooling hole down to x/D = 10. As compared to the cylindrical hole, both expanded holes show significantly improved thermal protection of the surface downstream of the ejection location, particularly at high blowing ratios. The laidback fanshaped hole provides a better lateral spreading of the ejected coolant than the fanshaped hole which leads to higher laterally averaged film-cooling effectiveness. Coolant passage crossflow Mach number and orientation strongly affect the flowfield of the jet being ejected from the hole and, therefore, have an important impact on film-cooling performance.

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Coolant Gas Injection on a Blunt-Nosed Re-Entry Vehicle

ASME 2013 Gas Turbine India Conference ◽

10.1115/gtindia2013-3738 ◽

2013 ◽

Author(s):

Jeswin Joseph ◽

S. R. Shine

Keyword(s):

Mach Number ◽

Film Cooling ◽

Thermal Protection ◽

Thermal Protection System ◽

Protection System ◽

Aerodynamic Heating ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

Injection Angle ◽

Mach Numbers

Very high thermal loads are expected in re-entry vehicles traveling at hypersonic Mach numbers due to severe aerodynamic heating. In the present study, numerical investigations are carried out to analyze the use of film cooling technology for a fully reusable and active thermal protection system of the re-entry vehicle. Simulations are done to examine the fundamental flow phenomenon and the performance of blunt body film cooling in hypersonic flows. Simulations are conducted for a blunt -nosed spacecraft flying at Mach numbers varying from 4 to 8 and 40 deg angle of attack. Film cooling holes are provided on the bottom of the blunt-nosed body. Standard values at an altitude of 30 km are used as in flow boundary conditions. The dependency of blowing ratios, stream-wise injection angle and inlet Mach number on the film cooling effectiveness are investigated. It is observed that the film cooling effectiveness reduces with increase in coolant injection angle. The film cooling performance is found to be decreasing with increase in Mach number. The results could provide useful inputs for optimization of an active thermal protection system of re-entry vehicles.

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Effectiveness of Normal and Angled Slot Cooling

Volume 3: Turbo Expo 2004 ◽

10.1115/gt2004-53248 ◽

2004 ◽

Author(s):

A. C. Smith ◽

J. H. Hatchett ◽

A. C. Nix ◽

W. F. Ng ◽

K. A. Thole ◽

...

Keyword(s):

Mach Number ◽

Film Cooling ◽

Fluid Dynamic ◽

Current Experiment ◽

Cfd Simulations ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

Blowing Ratio ◽

Mach Numbers ◽

Lift Off

An experimental and numerical investigation was conducted to determine the film cooling effectiveness of a normal slot and angled slot under realistic engine Mach number conditions. Freestream Mach numbers of 0.65 and 1.3 were tested. For the normal slot, hot gas ingestion into the slot was observed at low blowing ratios (M < 0.25). At high blowing ratios (M > 0.6) the cooling film was observed to “lift off” from the surface. For the 30° angled slot, the data was found to collapse using the blowing ratio as a scaling parameter. Results from the current experiment were compared with the subsonic data previously published. For the angle slot, at supersonic freestream Mach number, the current experiment shows that at the same x/Ms, the film-cooling effectiveness increases by as much as 25% as compared to the subsonic case. The results of the experiment also show that at the same x/Ms, the film cooling effectiveness of the angle slot is considerably higher than the normal slot, at both subsonic and supersonic Mach numbers. The flow physics for the slot tests considered here are also described with computational fluid dynamic (CFD) simulations in the subsonic and supersonic regimes.

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Conjugate Heat Transfer Analysis to Assess Overall Cooling Effectiveness of High Pressure Turbine Nozzle With Optimized Film Cooling Hole Arrangements

Volume 5A: Heat Transfer ◽

10.1115/gt2019-91194 ◽

2019 ◽

Author(s):

Young Seok Kang ◽

Dong-Ho Rhee ◽

Sanga Lee ◽

Bong Jun Cha

Keyword(s):

Heat Transfer ◽

High Pressure ◽

Film Cooling ◽

Conjugate Heat Transfer ◽

Heat Transfer Analysis ◽

Internal Cooling ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness ◽

Baseline Design ◽

Pressure Surface

Abstract Conjugate heat transfer analysis method has been highlighted for predicting heat exchange between fluid domain and solid domain inside high-pressure turbines, which are exposed to very harsh operating conditions. Then it is able to assess the overall cooling effectiveness considering both internal cooling and external film cooling at the cooled turbine design step. In this study, high-pressure turbine nozzles, which have three different film cooling holes arrangements, were numerically simulated with conjugate heat transfer analysis method for predicting overall cooling effectiveness. The film cooling holes distributed over the nozzle pressure surface were optimized by minimizing the peak temperature, temperature deviation. Additional internal cooling components such as pedestals and rectangular rib turbulators were modeled inside the cooling passages for more efficient heat transfer. The real engine conditions were given for boundary conditions to fluid and solid domains for conjugate heat transfer analysis. Hot combustion gas properties such as specific heat at constant pressure and other transport properties were given as functions of temperature. Also, the conductivity of Inconel 718 was also given as a function of temperature to solve the heat equation in the nozzle solid domain. Conjugate heat transfer analysis results showed that optimized designs showed better cooling performance, especially on the pressure surface due to proper staggering and spacing hole-rows compared to the baseline design. The overall cooling performances were offset from the adiabatic film cooling effectiveness. Locally concentrated heat transfer and corresponding high cooling effectiveness region appeared where internal cooling effects were overlapped in the optimized designs. Also, conjugate heat transfer analysis results for the optimized designs showed more uniform contours of the overall cooling effectiveness compared to the baseline design. By varying the coolant mass flow rate, it was observed that pressure surface was more sensitive to the coolant mass flow rate than nozzle leading edge stagnation region and suction surface. The CHT results showed that optimized designs to improve the adiabatic film cooling effectiveness also have better overall cooling effectiveness.

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