Design Optimization of Networks of Cooling Passages

2005 ◽  
Author(s):  
Nenad Jelisavcic ◽  
Thomas J. Martin ◽  
Ramon J. Moral ◽  
Debasis Sahoo ◽  
George S. Dulikravich ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Coolant Flow ◽  
Flow Rates ◽  
One Dimensional ◽  
Multi Objective ◽  
Internally Cooled

A one-dimensional thermo-fluid flow network analysis program COOLNET was developed to predict coolant flow rates, total coolant pressures, bulk coolant total temperatures, and internal heat transfer coefficient distributions, inside internally cooled objects. The coolant passages were allowed to be an arbitrary network of one-dimensional fluid elements or tubes. Geometric parameters of each passage were optimized using a hybrid multi-objective optimization algorithm. For the chosen Pratt & Whitney gas turbine the cross-sectional areas, hydraulic diameters, ribs’ heights and angles in each passage were optimized in order to minimize coolant flow rate, coolant total pressure loss, and total heat removed by coolant from the system. Validation of the hybrid multi-objective optimizer was performed with various test functions. Also, validation of the COOLNET was done with existing Pratt & Whitney gas turbine blade. Program OBJ was written to connect hybrid multi-objective optimizer and COOLNET. Optimization process was visualized using Tecplot (commercial software). Program PLOT was written to write input file for Tecplot, purpose of PLOT was to visualize initial and optimized results. Analysis for the best optimal result is given. The resulting COOLNET analysis provided coolant flow rates, total and static pressures and temperatures, and the heat transfer coefficient of each fluid element. The COOLNET analysis algorithm would typically converge in 50 iterations requiring about 5 seconds of CPU time on a 3.0 GHz processor. Results demonstrated that in case of an internally cooled gas turbine blade an improvement in overall performance is possible. A typical design optimization required between 500-1,000 iterations with a population of 30 designs. Thus, total number of configurations analyzed was approximately 15,000-30,000.

scholarly journals Heat Transfer and Pressure Distributions on a Gas Turbine Blade Tip

2000 ◽  
Vol 122 (4) ◽  
pp. 717-724 ◽  
Author(s):  
Gm. S. Azad ◽  
Je-Chin Han ◽  
Shuye Teng ◽  
Robert J. Boyle
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Turbulence Intensity ◽  
Rotor Blade ◽  
Static Pressure ◽  
Pressure Distributions ◽  
Blade Tip

Heat transfer coefficient and static pressure distributions are experimentally investigated on a gas turbine blade tip in a five-bladed stationary linear cascade. The blade is a two-dimensional model of a first-stage gas turbine rotor blade with a blade tip profile of a GE-E3 aircraft gas turbine engine rotor blade. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1×106. The middle 3-blade has a variable tip gap clearance. All measurements are made at three different tip gap clearances of about 1, 1.5, and 2.5 percent of the blade span. Heat transfer measurements are also made at two different turbulence intensity levels of 6.1 and 9.7 percent at the cascade inlet. Static pressure measurements are made in the midspan and the near-tip regions as well as on the shroud surface, opposite the blade tip surface. Detailed heat transfer coefficient distributions on the plane tip surface are measured using a transient liquid crystal technique. Results show various regions of high and low heat transfer coefficient on the tip surface. Tip clearance has a significant influence on local tip heat transfer coefficient distribution. Heat transfer coefficient also increases about 15–20 percent along the leakage flow path at higher turbulence intensity level of 9.7 over 6.1 percent. [S0889-504X(00)00404-9]

Effect of Squealer Geometry Arrangement on a Gas Turbine Blade Tip Heat Transfer

2002 ◽  
Vol 124 (3) ◽  
pp. 452-459 ◽  
Author(s):  
Gm Salam Azad ◽  
Je-Chin Han ◽  
Ronald S. Bunker ◽  
C. Pang Lee
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Gas Turbine Blade ◽  
Blade Tip

This study investigates the effect of a squealer tip geometry arrangement on heat transfer coefficient and static pressure distributions on a gas turbine blade tip in a five-bladed stationary linear cascade. A transient liquid crystal technique is used to obtain detailed heat transfer coefficient distribution. The test blade is a linear model of a tip section of the GE E3 high-pressure turbine first stage rotor blade. Six tip geometry cases are studied: (1) squealer on pressure side, (2) squealer on mid camber line, (3) squealer on suction side, (4) squealer on pressure and suction sides, (5) squealer on pressure side plus mid camber line, and (6) squealer on suction side plus mid camber line. The flow condition during the blowdown tests corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1×106. Results show that squealer geometry arrangement can change the leakage flow and results in different heat transfer coefficients to the blade tip. A squealer on suction side provides a better benefit compared to that on pressure side or mid camber line. A squealer on mid camber line performs better than that on a pressure side.

Effect of Squealer Geometry Arrangement on Gas Turbine Blade Tip Heat Transfer

2001 ◽  
Author(s):  
Gm Salam Azad ◽  
Je-Chin Han ◽  
Ronald S. Bunker ◽  
C. Pang Lee
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Gas Turbine Blade ◽  
Blade Tip

Abstract This study investigates the effect of a squealer tip geometry arrangement on heat transfer coefficient and static pressure distributions on a gas turbine blade tip in a five-bladed stationary linear cascade. A transient liquid crystal technique is used to obtain detailed heat transfer coefficient distribution. The test blade is a linear model of a tip section of the GE E3 high-pressure turbine first stage rotor blade. Six tip geometry cases are studied: 1) squealer on pressure side, 2) squealer on mid camber line, 3) squealer on suction side, 4) squealer on pressure and suction sides, 5) squealer on pressure side plus mid camber line, and 6) squealer on suction side plus mid camber line. The flow condition corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1 × 106. Results show that squealer geometry arrangement can change the leakage flow and results in different heat transfer coefficients to the blade tip. A squealer on suction side provides a better benefit compared to that on pressure side or mid camber line. A squealer on mid camber line performs better than that on a pressure side.

Effect of Rotation on a Gas Turbine Blade Internal Cooling System: Numerical Investigation

2016 ◽  
Author(s):  
E. Burberi ◽  
D. Massini ◽  
L. Cocchi ◽  
L. Mazzei ◽  
A. Andreini ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Flow Field ◽  
Transfer Coefficient ◽  
Numerical Investigation ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Cooling System ◽  
Leading Edge ◽  
Internal Cooling

Increasing turbine inlet temperature is one of the main strategies used to accomplish the demands of increased performance of modern gas turbines. As a consequence, optimization of the cooling system is of paramount importance in gas turbine development. Leading edge represents a critical part of cooled nozzles and blades, given the presence of the hot gases stagnation point and the unfavourable geometry for cooling. This paper reports the results of a numerical investigation aimed at assessing the rotation effects on the heat transfer distribution in a realistic leading edge internal cooling system of a high pressure gas turbine blade. The numerical investigation was carried out in order to support and to allow an in-depth understanding of the results obtained in a parallel experimental campaign. The model is composed of a trapezoidal feeding channel which provides air to the cold bridge system by means of three large racetrack-shaped holes, generating coolant impingement on the internal concave leading edge surface, whereas four big fins assure the jets confinement. Air is then extracted through 4 rows of 6 holes reproducing the external cooling system composed of shower-head and film cooling holes. Experiments were performed in static and rotating conditions replicating the typical range of jet Reynolds number (Rej) from 10000 to 40000 and Rotation number (Roj) up to 0.05, for three crossflow cases representative of the working condition that can be found at blade tip, midspan and hub, respectively. Experimental results in terms of flow field measurements on several internal planes and heat transfer coefficient on the LE internal surface have been performed on two analogous experimental campaigns at University of Udine and University of Florence respectively. Hybrid RANS-LES models were used for the simulations, such as Scale Adaptive Simulation (SAS) and Detached Eddy Simulation (DES), given their ability to resolve the complex flow field associated with jet impingement. Numerical flow field results are reported in terms of both jet velocity profiles and 2D vector plots on symmetry and transversal internal planes, while the heat transfer coefficient distributions are presented as detailed 2D maps together with radial and tangential averaged Nusselt number profiles. A fairly good agreement with experimental measurements is observed, which represent a validation of the adopted computational model. As a consequence, the computed aerodynamic and thermal fields also allow an in-depth interpretation of the experimental results.

Numerical Investigation of Heat Transfer Enhancement in Gas Turbine Blade by Air Film Cooling and Different Types of Ribs

2013 ◽  
Author(s):  
Maryam Pourhasanzadeh
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Convection Heat Transfer ◽  
Convection Heat ◽  
Convection Heat Transfer Coefficient ◽  
The Heat Transfer Coefficient

In this research, numerical studies have been carried out for a film cooling jet on a gas turbine blade in the presence of different kinds of ribs. Simulations are performed for a film cooling jet inclined at 30-degrees. The effect of coolant jet velocity on convection heat transfer coefficient (h) is investigated. In addition, various ribs geometries and their distance from the blade surface are examined. It is shown that a combination of the rib and the film cooling jet stimulate the momentum and thermal boundary layers and subsequently improve the convection heat transfer coefficient. It is indicated that the heat transfer coefficient is dependent on the height of the rib and there is an inverse relation with the rib distance from the plate. Moreover, an increase in coolant jet velocity causes the increase of the heat transfer coefficient. The results show a significant improvement of the heat transfer coefficient over three times more than h on a blade without any ribs or coolant jet.

scholarly journals Experimental and Numerical Investigation of Effect of Blowing Ratio on Film Cooling Effectiveness and Heat Transfer Coefficient Over a Gas Turbine Blade Leading Edge Film Cooling Configurations

2013 ◽  
Author(s):  
Yepuri Giridhara Babu ◽  
Gururaj Lalgi ◽  
Ashok Babu Talanki Puttarangasetty ◽  
Jesuraj Felix ◽  
Sreenivas Rao V. Kenkere ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbine Blade ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

Film cooling is one of the cooling techniques to cool the hot section components of a gas turbine engines. The gas turbine blade leading edges are the vital parts in the turbines as they are directly hit by the hot gases, hence the optimized cooling of gas turbine blade surfaces is essential. This study aims at investigating the film cooling effectiveness and heat transfer coefficient experimentally and numerically for the three different gas turbine blade leading edge models each having the one row of film cooling holes at 15, 30 and 45 degrees hole orientation angle respectively from stagnation line. Each row has the five holes with the hole diameter of 3mm, pitch of 20mm and has the hole inclination angle of 20deg. in spanwise direction. Experiments are carried out using the subsonic cascade tunnel facility of National Aerospace Laboratories, Bangalore at a nominal flow Reynolds number of 1,00,000 based on the leading edge diameter, varying the blowing ratios of 1.2, 1.50, 1.75 and 2.0. In addition, an attempt has been made for the film cooling effectiveness using CFD simulation, using k-€ realizable turbulence model to solve the flow field. Among the considered 15, 30 and 45 deg. models, both the cooling effectiveness and heat transfer coefficient shown the increase with the increase in hole orientation angle from stagnation line. The film cooling effectiveness increases with the increase in blowing ratio upto 1.5 for the 15 and 30 deg. models, whereas on the 45 deg. model the increase in effectiveness shown upto the blowing ratio of 1.75. The heat transfer coefficient values showed the increase with the increase in blowing ratio for all the considered three models. The CFD results in the form of temperature, velocity contours and film cooling effectiveness values have shown the meaningful results with the experimental values.

Heat Transfer Coefficient on the Squealer Tip and Near Squealer Tip Regions of a Gas Turbine Blade

2002 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Rotor Blade ◽  
Intensity Level ◽  
Gas Turbine Blade ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients

Detailed heat transfer coefficient distributions on a squealer tip of a gas turbine blade were measured using a hue detection based transient liquid crystals technique. The heat transfer coefficients on the shroud and near tip region of the pressure and suction sides of a blade were also measured. Tests were performed on a five-bladed linear cascade with blow down facility. The blade was a 2-dimensional model of a first stage gas turbine rotor blade with a profile of a GE-E3 aircraft gas turbine engine rotor blade. The Reynolds number based on the cascade exit velocity and axial chord length of a blade was 1.1×106 and the total turning angle of the blade was 97.7°. The overall pressure ratio was 1.23 and the inlet and exit Mach number were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. The heat transfer measurements were taken at the three different tip gap clearances of 1.0%, 1.5% and 2.5% of blade span. Results showed that the overall heat transfer coefficient on the squealer tip was higher than that on the shroud and the near tip region of the pressure and suction side. Results also showed that the heat transfer coefficients on the squealer tip and its shroud were lower than that on the plane tip and shroud, but the heat transfer coefficients on the near tip region of suction and pressure sides were higher for the squealer tip case.

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