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.

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]

scholarly journals Full Coverage Discrete Hole Wall Cooling: Cooling Effectiveness

1984 ◽  
Author(s):  
G. E. Andrews ◽  
M. L. Gupta ◽  
M. C. Mkpadi
Keyword(s):  
Heat Transfer ◽  
Low Temperature ◽  
Film Cooling ◽  
Test Facility ◽  
Coolant Temperature ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Full Coverage ◽  
Hole Wall

The development of a test facility for investigating full coverage discrete hole wall cooling for gas turbine combustion chamber wall cooling is described. A low temperature test condition of 750K mainstream temperature and 300K coolant temperature was used to investigate the influence of coolant flow rate at a constant cross flow Mach number. Practical combustion conditions of 2100K combustor temperature and 700K coolant temperature are investigated to establish the validity of applying the low temperature results to practical conditions. For both situations a heat balance programme, taking into account the heat transfer within the wall was used to compute the film heat transfer coefficients. The mixing of the coolant air with the mainstream gases was studied through boundary layer temperature and CO2 profiles. It was shown that entrainment of hot flame gases between the injection holes resulted in a very low ‘adiabatic’ film cooling effectiveness.

scholarly journals Heat Transfer Phenomena in Gas Turbines

1980 ◽  
Author(s):  
J. R. Taylor
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Turbine Blades ◽  
Gas Turbine Engine ◽  
Transfer Coefficients ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients

A discussion of the problems encountered in prediction of heat transfer in the turbine section of a gas turbine engine is presented. Areas of current gas turbine engine is presented. Areas of current concern to designers where knowledge is deficient or lacking are elucidated. Consideration is given to methods and problems associated with determination of heat transfer coefficients, external gas temperatures, and, where applicable, film cooling effectiveness. The paper is divided into parts dealing with turbine airfoil heat transfer, endwall heat transfer, and heat transfer in the internal cavities of cooled turbine blades. Recent literature dealing with these topics is listed.

Film Cooling Effectiveness and Mass/Heat Transfer Coefficient Downstream of One Row of Discrete Holes

1999 ◽  
Vol 121 (2) ◽  
pp. 225-232 ◽  
Author(s):  
R. J. Goldstein ◽  
P. Jin ◽  
R. L. Olson
Keyword(s):  
Heat Transfer ◽  
Mass Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness

A special naphthalene sublimation technique is used to study the film cooling performance downstream of one row of holes of 35 deg inclination angle with 3d hole spacing and relatively small hole length to diameter ratio (L/d = 6.3). Both film cooling effectiveness and mass/heat transfer coefficient are determined for blowing rates from 0.5 to 2.0 with density ratio of 1.0. The mass transfer coefficient is measured using pure air film injection, while the film cooling effectiveness is derived from comparison of mass transfer coefficients obtained following injection of naphthalene-vapor-saturated air with those from pure air injection. This technique enables one to obtain detailed local information on film cooling performance. The laterally averaged and local film cooling effectiveness agree with previous experiments. The difference between mass/heat transfer coefficients and previous heat transfer results indicates that conduction error may play an important role in the earlier heat transfer measurements.

Fan Shaped and Cylindrical Holes Studied in Vane Film Cooling Test Rig

2010 ◽  
Author(s):  
Mats Kinell ◽  
Esa Utriainen ◽  
Jonas Hyle´n ◽  
Jonas Gustavsson ◽  
Andreas Bradley ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Coefficient ◽  
Suction Side ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Coefficient Distribution ◽  
Shaped Hole

In order to optimize the vane film cooling and thereby increase the efficiency of a gas turbine, different film cooling configurations were experimentally investigated. Dynamic similarity was obtained regarding main flow Reynolds number, airfoil pressure coefficient, adiabatic wall temperature and film cooling ejection ratio. The maximum reached Mach number was 0.52. The geometry of the test section, consisting of one vane and two flow paths, was modified in order to meet the dimensionless pressure coefficient distribution around the airfoil experienced by a full stage airfoil. This would ascertain that scaled but engine realistic pressure gradients would be achieved in the rig test. During the test, the cold airfoil was suddenly imposed to a hot main stream and the evaluation of both the film cooling effectiveness and the heat transfer coefficient distribution on the visiable surface could be done at one single test using time-resolved temperature measurements obtained through IR thermography. A high resolution MWIR camera was used together with a silicon viewing window. The post-processing allowed for corrections regarding emissions and determination of the desired parameters on the vane surface. Results, heat transfer coefficients and film cooling effectiveness, for fan shaped and cylindrical film cooling holes configurations are compared. The results show clear benefit of using shaped holes over cylindrical ditto, especially on the suction side where near hole film effectiveness is enhanced by approximately 25%, but the results also show that this benefit diminishes to nothing in the suction side trailing edge region. The local heat transfer coefficients are generally lower for the shaped hole configurations. Contrary to the film effectiveness the shaped holes configurations show lower heat transfer coefficients also at the suction side trailing edge region, making use of the shaped hole configurations superior to cylindrical ones as the heat flux to the surface is reduced. Numerical predictions using a boundary layer code, TEXSTAN, and CFD, for a smooth wall configuration corresponds well with the measured results.

Effect of Showerhead Injection on Superposition of Multi-Row Pressure Side Film Cooling With Fan Shaped Holes

2003 ◽  
Author(s):  
Martin Schneider ◽  
Sacha Parneix ◽  
Jens von Wolfersdorf
Keyword(s):  
Boundary Layer ◽  
Film Cooling ◽  
Density Ratio ◽  
Test Facility ◽  
Main Stream ◽  
Test Model ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Single Row

In the present study the effect of approach boundary layer conditions on film cooling effectiveness superposition for pressure side fan shaped holes on a first stage vane is investigated. More particularly, the effect of showerhead cooling on the film effectiveness of downstream pressure side rows is addressed. The experimental test facility used is a continuously running , two passage, linear cascade wind tunnel equipped with a central vane and contoured side walls. Main stream stagnation conditions and pressure measurements on the vane external surface were taken to determine the isentropic Mach number distribution. The turbulence level generated with a bar grid is around 15%. Film cooling effectiveness has been determined with the narrow banded thermochromic liquid crystal steady state technique. Mainstream as well as coolant flow could be heated to shift the iso-temperature contours across the vane. Carbon dioxide was used as coolant gas to better match the density ratio in the experimental facility to engine conditions. The test model is a research vane equipped with four rows of cylindrical holes showerhead and three rows of fan shaped holes along the pressure side at typical inclination angles to the surface. The study incorporates single row blowing with the following row approach conditions: 1) no showerhead injection, 2) boundary layer trip, 3) isothermal showerhead blowing, and multi-row blowing with and without showerhead blowing. The results are used to investigate the applicability of the single row results superposition approach for multiple-row injection. Simulating the appropriate aerodynamic conditions during individual row measurements improves the superposition prediction in comparison to the multi-row results.

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.

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

2003 ◽  
Vol 125 (4) ◽  
pp. 648-657 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a two-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film cooling holes on the pressure side near the tip region and the tip surface along the camber line. Hue detection based transient liquid crystals technique was used to measure heat transfer coefficients and film cooling effectiveness. All measurements were done for the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1×106 and the total turning angle of the blade was 97.9 deg. The overall pressure ratio was 1.2 and the inlet and exit Mach numbers were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

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.

scholarly journals Numerical Prediction of Film Cooling Effectiveness and the Associated Aerodynamic Losses With a Three-Dimensional Calculation Procedure

1990 ◽  
Author(s):  
G. H. Dibelius ◽  
R. Pitt ◽  
B. Wen
Keyword(s):  
Experimental Data ◽  
Film Cooling ◽  
Reverse Flow ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Temperature Pattern ◽  
Main Stream ◽  
Cooling Efficiency ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Losses

Film cooling of turbine blades by injecting air through holes or slots affects the main stream flow. A numerical model has been developed to predict the resulting three-dimensional flow and the temperature pattern under steady flow conditions. An elliptic procedure is used in the near injection area to include reverse flow situations, while in the upstream area as well as far downstream a partial-parabolic procedure is applied. As first step an adiabatic wall has been assumed as boundary condition, since for this case experimental data are readily available for comparison. At elevated momentum blowing rates, zones of reverse flow occur downstream of the injection holes resulting in a decrease of cooling efficiency. A variation of the relevant parameters momentum blowing rate m, injection angle α and ratio of hole spacing to diameter s/d revealed the combination of m ≈ 1, α ≈ 30° and s/d ≈ 2 to be the optimum with respect to the averaged cooling efficiency and to the aerodynamic losses. Cooling is more efficient with slots than with a row of holes not considering the related problems of manufacture and service life. The calculated temperature patterns compare well with the experimental data available.

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