A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall: Part II — Heat Transfer and Film Cooling Measurements

2015 ◽  
Author(s):  
Johannes Kneer ◽  
Franz Puetz ◽  
Achmed Schulz ◽  
Hans-Jörg Bauer
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Superposition Principle ◽  
Density Ratio ◽  
Open Loop ◽  
Test Facility ◽  
Gas Temperature ◽  
Test Rig ◽  
Linear Cascade ◽  
Exit Mach Number

The present work is part of a comprehensive heat transfer and film-cooling study on a locally cooled non-axisymmetric contoured turbine endwall. A new test rig consisting of a linear cascade of three prismatic vanes at unity scale and exchangeable endwall has been established. The rig is operated in an open-loop configuration at a reduced main gas temperature of 425 K, an exit Mach number of 0.5 and an exit Reynolds number of 1.6×106. Air is used both as main gas and coolant; a realistic density ratio is achieved by cooling the coolant below freezing. In Part I [1] of the study aerodynamic measurements are presented. This paper concentrates on film-cooling of the contoured endwall with special emphasis on data acquisition and reduction for the application of the superposition principle of film cooling. The first experimental results from thermographic measurements are discussed.

A New Test Facility to Investigate Film Cooling on a Nonaxisymmetric Contoured Turbine Endwall—Part II: Heat Transfer and Film Cooling Measurements

2016 ◽  
Vol 138 (7) ◽  
Author(s):  
Johannes Kneer ◽  
Franz Puetz ◽  
Achmed Schulz ◽  
Hans-Jörg Bauer
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Superposition Principle ◽  
Density Ratio ◽  
Open Loop ◽  
Test Facility ◽  
Gas Temperature ◽  
Test Rig ◽  
Linear Cascade ◽  
Exit Mach Number

The present work is part of a comprehensive heat transfer and film-cooling study on a locally cooled nonaxisymmetric contoured turbine endwall. A new test rig consisting of a linear cascade of three prismatic vanes at unity scale and exchangeable endwall has been established. The rig is operated in an open-loop configuration at a reduced main gas temperature of 425 K, an exit Mach number of 0.5, and an exit Reynolds number of 1.6 × 106. Air is used both as main gas and coolant; a realistic density ratio is achieved by cooling the coolant below freezing. In the first part of the study, aerodynamic measurements are presented. This paper concentrates on film cooling of the contoured endwall with special emphasis on data acquisition and reduction for the application of the superposition principle of film cooling. The first experimental results from thermographic measurements are discussed.

Endwall Heat Transfer and Cooling Performance of a Transonic Turbine Vane with Upstream Injection Flow

2021 ◽  
pp. 1-24
Author(s):  
Zhigang LI ◽  
Bo Bai ◽  
Jun Li ◽  
Shuo Mao ◽  
Wing Ng ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Density Ratio ◽  
Gas Temperature ◽  
Blow Down ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Injection Flow ◽  
Coolant Injection ◽  
Exit Mach Number

Abstract Detailed experimental and numerical studies on endwall heat transfer and cooling performance with coolant injection flow through upstream discrete holes is presented in this paper. High resolution heat transfer coefficient (HTC) and adiabatic film cooling effectiveness values were measured using a transient infrared thermography technique on an axisymmetric contoured endwall. The tests were performed in a transonic linear cascade blow-down wind tunnel facility. Conditions were representative of a land-based power generation turbine with exit Mach number of 0.85 corresponding to exit Reynolds number of 1.5 × 106, based on exit condition and axial chord length. A high turbulence level of 16% with an integral length scale of 3.6%P was generated using inlet turbulence grid to reproduce the typical turbulence conditions in real turbine. Low temperature air was used to simulate the typical coolant-to-mainstream condition by controlling two parameters of the upstream coolant injection flow: mass flow rate to determine the coolant-to-mainstream blowing ratio (BR = 2.5, 3.5), and gas temperature to determine the density ratio (DR = 1.2). To highlight the interactions between the upstream coolant flow and the passage secondary flow combined with the influence on the endwall heat transfer and cooling performance, a comparison of CFD predictions to experimental results was performed by solving steady-state Reynolds-Averaged Navier-Stokes (RANS) using the commercial CFD solver ANSYS Fluent V.15.

Experimental Study of Effusion Cooling With Pressure-Sensitive Paint

2017 ◽  
Vol 139 (5) ◽  
Author(s):  
Guanghua Wang ◽  
Gustavo Ledezma ◽  
James DeLancey ◽  
Anquan Wang
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Cfd Simulation ◽  
Density Ratio ◽  
Open Loop ◽  
Pressure Sensitive Paint ◽  
Gas Temperature ◽  
Eddy Simulation ◽  
Pressure Sensitive ◽  
The Impact

Gas turbines overall efficiency enhancement requires further increasing of the firing temperature and decreasing of cooling flow usage. Multihole (or effusion, or full-coverage) film cooling is widely used for hot gas path components cooling in modern gas turbines. The present study focused on the adiabatic film effectiveness measurement of a round multihole flat-plate coupon. The measurements were conducted in a subsonic open-loop wind tunnel with a generic setup to cover different running conditions. The test conditions were characterized by a constant main flow Mach number of 0.1 with constant gas temperature. Adiabatic film effectiveness was measured by pressure-sensitive paint (PSP) through mass transfer analogy. CO2 was used as the coolant to reach the density ratio of 1.5. Rig computational fluid dynamics (CFD) simulation was conducted to evaluate the impact of inlet boundary layer on testing. Experimental data cover blowing ratios (BRs) at 0.4, 0.6, 0.8, 1.0, and 2.0. Both 2D maps and lateral average profiles clearly indicated that the film effectiveness increases with increasing BR for BR < 0.8 and decreases with increasing BR for BR > 0.8. This observation agreed with coolant jet behavior of single film row, i.e., attached, detached then reattached, and fully detached. PSP data quality was then discussed in detail for validating large eddy simulation.

scholarly journals Experimental Study of Showerhead Cooling on a Cylinder Comparing Several Configurations Using Cylindrical and Shaped Holes

1999 ◽  
Vol 122 (1) ◽  
pp. 161-169 ◽  
Author(s):  
H. Reiss ◽  
A. Bo¨lcs
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Performance Enhancement ◽  
Turbine Blades ◽  
Density Ratio ◽  
Test Facility ◽  
Thermochromic Liquid Crystal ◽  
Transfer Coefficients ◽  
Film Cooling Effectiveness ◽  
Shaped Hole

Film cooling and heat transfer measurements on a cylinder model have been conducted using the transient thermochromic liquid crystal technique. Three showerhead cooling configurations adapted to leading edge film cooling of gas turbine blades were directly compared: “classical” cylindrical holes versus two types of shaped hole exits. The experiments were carried out in a free jet test facility at two different flow conditions, Mach numbers M=0.14 and M=0.26, yielding Reynolds numbers based on the cylinder diameter of 8.6e4 and 1.55e5, respectively. All experiments were done at a mainstream turbulence level of Tu=7 percent with an integral length scale of Lx=9.1 mmM=0.14, or Lx=10.5 mmM=0.26, respectively. Foreign gas injection CO2 was used, yielding an engine-near density ratio of 1.6, with blowing ratios ranging from 0.6 to 1.5. Detailed experimental results are shown, including surface distributions of film cooling effectiveness and local heat transfer coefficients. Additionally, heat transfer and heat load augmentation due to injection with respect to the uncooled cylinder are reported. For a given cooling gas consumption, the laid-back shaped hole exits lead to a clear enhancement of the cooling performance compared to cylindrical exits, whereas laterally expanded holes give only slight performance enhancement. [S0889-504X(00)01801-8]

Experimental and Numerical Investigations of Effusion Cooling for High Pressure Turbine Components: Part 1 — Experimental Study With PSP

2016 ◽  
Author(s):  
Guanghua Wang ◽  
Gustavo Ledezma ◽  
James DeLancey ◽  
Anquan Wang
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Cfd Simulation ◽  
Density Ratio ◽  
Open Loop ◽  
Gas Temperature ◽  
Cooling Flow ◽  
Eddy Simulation ◽  
Jet Behavior ◽  
Effectiveness Measurement

Gas turbines overall efficiency enhancement requires further increasing of the firing temperature and decreasing of cooling flow usage. Multi-Hole (or effusion, full coverage) film cooling is widely used for hot gas path components cooling in modern gas turbines. The present study focused on the adiabatic film effectiveness measurement of a round multi-hole flat plate coupon. The measurements were conducted in a subsonic open loop wind tunnel with a generic setup to cover different running conditions. The test conditions were characterized by a constant main flow Mach number of 0.1 with constant gas temperature. Adiabatic film effectiveness was measured by Pressure Sensitive Paint (PSP) through mass transfer analogy. CO2 was used as the coolant to reach the density ratio of 1.5. Rig CFD simulation was conducted to evaluate impact of inlet boundary layer on testing. Experimental data covers blowing ratios (BR) at 0.4, 0.6, 0.8, 1.0 and 2.0. Both 2D maps and lateral average profiles clearly indicated that film effectiveness increases with increasing BR for BR<0.8 and decreases with increasing BR for BR>0.8. This observation agreed with coolant jet behavior of single film row, i.e. attached, detached then reattached, and fully detached. PSP data quality was then discussed in detail for validating Large Eddy Simulation in Part 2.

scholarly journals Experimental Study of Showerhead Cooling on a Cylinder Comparing Several Configurations Using Cylindrical and Shaped Holes

1999 ◽  
Author(s):  
H. Reiss ◽  
A. Bölcs
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Turbine Blades ◽  
Density Ratio ◽  
Test Facility ◽  
Main Stream ◽  
Transfer Coefficients ◽  
Surface Distribution ◽  
Film Cooling Effectiveness ◽  
Shaped Hole

Film cooling and heat transfer measurements on a cylinder model have been conducted using the transient thermochromic liquid crystal technique. Three showerhead cooling configurations adapted to leading edge film cooling of gas turbine blades were directly compared: ‘classical’ cylindrical holes versus two types of shaped hole exits. The experiments were carried out in a free jet test facility at two different flow conditions, Mach numbers M = 0.14 and M = 0.26, yielding Reynolds numbers based on the cylinder diameter of 8.6e4 and 1.55e5, respectively. All experiments were done at a main stream turbulence level of Tu = 7% with an integral lengthscale of Lx = 9.1mm (M = 0.14), or Lx = 10.5mm (M = 0.26) respectively. Foreign gas injection (CO2) was used yielding an engine-near density ratio of 1.6, with blowing ratios ranging from 0.6 to 1.5. Detailed experimental results are shown, including surface distribution of film cooling effectiveness and local heat transfer coefficients. Additionally, heat transfer and heat load augmentation due to injection with respect to the uncooled cylinder are reported. For a given cooling gas consumption the laid-back shaped hole exits lead to a clear enhancement of the cooling performance compared to cylindrical exits, whereas laterally expanded holes give only slight performance enhancement.

Conjugate Heat Transfer and Film Cooling of a Multi-Stage Cooling Scheme

2008 ◽  
Author(s):  
M. Ghorab ◽  
S. I. Kim ◽  
I. Hassan
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Gas Turbines ◽  
Conjugate Heat Transfer ◽  
Density Ratio ◽  
Design Parameters ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Multi Stage ◽  
Internal Gap

Cooling techniques play a key role in improving efficiency and power output of modern gas turbines. The conjugate technique of film and impingement cooling schemes is considered in this study. The Multi-Stage Cooling Scheme (MSCS) involves coolant passing from inside to outside turbine blade through two stages. The first stage; the coolant passes through first hole to internal gap where the impinging jet cools the external layer of the blade. Finally, the coolant passes through the internal gap to the second hole which has specific designed geometry for external film cooling. The effect of design parameters, such as, offset distance between two-stage holes, gap height, and inclination angle of the first hole, on upstream conjugate heat transfer rate and downstream film cooling effectiveness performance are investigated computationally. An Inconel 617 alloy with variable properties is selected for the solid material. The conjugate heat transfer and film cooling characteristics of MSCS are analyzed across blowing ratios of Br = 1 and 2 for density ratio, 2. This study presents upstream wall temperature distributions due to conjugate heat transfer for different gap design parameters. The maximum film cooling effectiveness with upstream conjugate heat transfer is less than adiabatic film cooling effectiveness by 24–34%. However, the full coverage of cooling effectiveness in spanwise direction can be obtained using internal cooling with conjugate heat transfer, whereas adiabatic film cooling effectiveness has narrow distribution.

scholarly journals Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

2012 ◽  
Vol 2012 ◽  
pp. 1-28 ◽  
Author(s):  
Phil Ligrani
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
High Speed ◽  
Density Ratio ◽  
Aerodynamic Loss ◽  
Pressure Losses ◽  
Loss Coefficients ◽  
Exit Mach Number

The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i) symmetric airfoils with no film cooling, (ii) symmetric airfoils with film cooling, (iii) cambered vanes with no film cooling, and (iv) cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.

Influence of Coolant Density on Turbine Platform Film-Cooling With Stator–Rotor Purge Flow and Compound-Angle Holes

2014 ◽  
Vol 6 (4) ◽  
Author(s):  
Kevin Liu ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Fundamental Parameters ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Exit Mach Number ◽  
Compound Angle

A detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform. The platform was cooled by purge flow from a simulated stator–rotor seal combined with discrete hole film-cooling. The cylindrical holes and laidback fan-shaped holes were accessed in terms of film-cooling effectiveness. This paper focuses on the effect of coolant-to-mainstream density ratio on platform film-cooling (DR = 1 to 2). Other fundamental parameters were also examined in this study—a fixed purge flow of 0.5%, three discrete-hole film-cooling blowing ratios between 1.0 and 2.0, and two freestream turbulence intensities of 4.2% and 10.5%. Experiments were done in a five-blade linear cascade with inlet and exit Mach number of 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 750,000 and was based on the exit velocity and chord length of the blade. The measurement technique adopted was the conduction-free pressure sensitive paint (PSP) technique. Results indicated that with the same density ratio, shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The optimum blowing ratio of 1.5 exists for the cylindrical holes, whereas the effectiveness for the shaped holes increases with an increase of blowing ratio. Results also indicate that the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity.

An Experimental Study of Turbine Vane Heat Transfer With Leading Edge and Downstream Film Cooling

1990 ◽  
Vol 112 (3) ◽  
pp. 477-487 ◽  
Author(s):  
N. V. Nirmalan ◽  
L. D. Hylton
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Leading Edge ◽  
External Heat ◽  
Gas Temperature ◽  
Temperature Ratio ◽  
External Heat Transfer ◽  
Turbine Vane

This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine vane external heat transfer. Steady-state experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters—Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio—were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The vane external heat transfer data obtained in this program indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil. The data presented in this paper illustrate the interaction of these variables and should provide the airfoil designer and computational analyst with the information required to improve heat transfer design capabilities for film-cooled turbine airfoils.

Sign in / Sign up

Export Citation Format

Share Document