Effect of Tip Gap and Squealer Geometry on Detailed Heat Transfer Measurements Over a High Pressure Turbine Rotor Blade Tip

2004 ◽  
Vol 126 (2) ◽  
pp. 221-228 ◽  
Author(s):  
Hasan Nasir ◽  
Srinath V. Ekkad ◽  
David M. Kontrovitz ◽  
Ronald S. Bunker ◽  
Chander Prakash
Keyword(s):  
Heat Transfer ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Transfer Coefficients ◽  
Blow Down ◽  
Turbine Rotor Blade ◽  
Heat Transfer Coefficients ◽  
Gap Height

The present study explores the effects of gap height and tip geometry on heat transfer distribution over the tip surface of a HPT first-stage rotor blade. The pressure ratio (inlet total pressure to exit static pressure for the cascade) used was 1.2, and the experiments were run in a blow-down test rig with a four-blade linear cascade. A transient liquid crystal technique was used to obtain the tip heat transfer distributions. Pressure measurements were made on the blade surface and on the shroud for different tip geometries and tip gaps to characterize the leakage flow and understand the heat transfer distributions. Two different tip gap-to-blade span ratios of 1% and 2.6% are investigated for a plane tip, and a deep squealer with depth-to-blade span ratio of 0.0416. For a shallow squealer with depth-to-blade span ratio of 0.0104, only 1% gap-to-span ratio is considered. The presence of the squealer alters the tip gap flow field significantly and produces lower overall heat transfer coefficients. The effects of different partial squealer arrangements are also investigated for the shallow squealer depth. These simulate partial burning off of the squealer in real turbine blades. Results show that some partial burning of squealers may be beneficial in terms of overall reduction in heat transfer coefficients over the tip surface.

Effect of Tip Gap and Squealer Geometry on Measured Heat Transfer Over a HPT Rotor Blade Tip

2003 ◽  
Author(s):  
Hasan Nasir ◽  
Srinath V. Ekkad ◽  
David M. Kontrovitz ◽  
Ronald S. Bunker ◽  
Chander Prakash
Keyword(s):  
Heat Transfer ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Transfer Coefficients ◽  
Blow Down ◽  
Pressure Surface ◽  
Heat Transfer Coefficients ◽  
Gap Flow ◽  
Gap Height

The present study explores the effects of gap height and tip geometry on heat transfer distribution over the tip surface of a HPT first stage rotor blade. The pressure ratio (inlet total pressure to exit static pressure for the cascade) used was 1.2, and the experiments were run in a blow-down test rig with a four-blade linear cascade. A transient liquid crystal technique was used to obtain the tip heat transfer distributions. Pressure measurements were made on the blade surface and on the shroud for different tip geometries and tip gaps to characterize the leakage flow and understand the heat transfer distributions. Two different tip gap-to-blade span ratio of 1% and 2.6% are investigated for a plane tip and a deep squealer with depth-to-blade span ratio of 0.0416. For a shallow squealer with depth-to-blade span ratio of 0.0104, only 1% gap-to-span ratio is considered. The presence of the squealer alters the tip gap flow field significantly and produces lower overall heat transfer coefficients. The effects of different partial squealer arrangements are also investigated for the shallow squealer depth. These simulate partial burning off of the squealer in real turbine blades. Results show that in some cases, partial burning of squealers along the pressure surface may be beneficial in terms of overall reduction in heat transfer coefficients over the tip surface compared to the plain tip.

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

2003 ◽  
Vol 125 (4) ◽  
pp. 669-677 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Intensity Level ◽  
Gas Turbine Blade ◽  
Transfer Coefficients ◽  
Blow Down ◽  
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 regions of the pressure and suction sides of a blade were also measured. Tests were performed on a five-bladed linear cascade with a blow-down facility. The blade was a two-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 deg. The overall pressure ratio was 1.2 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 percent. The heat transfer measurements were taken at the three different tip gap clearances of 1.0 percent, 1.5 percent, and 2.5 percent of blade span. Results showed that the overall heat transfer coefficients on the squealer tip were higher than that on the shroud surface and the near tip regions of the pressure and suction sides. 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. However, the reductions of heat transfer coefficients near the tip regions of the pressure and suction sides were not remarkable.

Comparison of Time-Resolved Turbine Rotor Blade Heat Transfer Measurements and Numerical Calculations

1991 ◽  
Author(s):  
R. S. Abhari ◽  
G. R. Guenette ◽  
A. H. Epstein ◽  
M. B. Giles
Keyword(s):  
Heat Transfer ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Time History ◽  
Turbine Rotor ◽  
Numerical Calculations ◽  
Test Facility ◽  
Time Resolved ◽  
Turbine Rotor Blade ◽  
History Of

Time-resolved turbine rotor blade heat transfer data are compared with ab initio numerical calculations. The data was taken on a transonic, 4-to-1 pressure ratio, uncooled, single-stage turbine in a short duration turbine test facility. The data consists of the time history of the heat transfer distribution about the rotor chord at midspan. The numerical calculation is a time accurate, 2-D, thin shear layer, multiblade row code known as UNSFLO. UNSFLO uses Ni’s Lax-Wendroff algorithm, conservative boundary conditions, and a time tilting algorithm to facilitate the calculation of the flow in multiple blade rows of arbitrary pitch ratio with relatively little computer time. The version used for this work had a simple algebraic Baldwin-Lomax turbulence model. The code is shown to do a good job of predicting the quantitative time history of the heat flux distribution. The wake/boundary layer and transonic interaction regions for suction and pressure surfaces are identified and the shortcomings of the current algebraic turbulence modelling in the code are discussed. The influence of hardware manufacturing tolerance on rotor heat transfer variation is discussed. A physical reasoning explaining the discrepancies between the unsteady measurement and the calculations for both the suction and pressure surfaces are given, which may be of use in improving future calculations and design procedures.

Experimental and Numerical Analysis of Gas Turbine Blades With Different Internal Cooling Geometries

2009 ◽  
Author(s):  
M. Eifel ◽  
V. Caspary ◽  
H. Ho¨nen ◽  
P. Jeschke
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Base Flow ◽  
Internal Cooling ◽  
Turbine Rotor Blade ◽  
Flow Phenomena ◽  
Sst Turbulence Model

This paper presents the effects of major geometrical modifications to the interior of a convection cooled gas turbine rotor blade. The analysis of the flow is performed experimentally with flow visualization via paint injection into water whereas the flow and the heat transfer are investigated numerically with Ansys CFX utilizing the SST turbulence model. Two sets of calculations are carried out, one under the same conditions as the experiments and another according to realistic hot gas conditions with conjugate heat transfer. The aim is to identify flow phenomena altering the heat transfer in the blade and to manipulate them in order to reduce the thermal load of the material. The operating point of the geometric base configuration is set to Re = 50,000 at the inlet while for the modified geometries the pressure ratio is held constant compared to the base. Flow structures and heat transfer conditions are evaluated and are linked to specific geometric features. Among several investigated configurations one could be identified that leads to a cooling effectiveness 15% larger compared to the base.

Experimental and Numerical Analysis of Gas Turbine Blades With Different Internal Cooling Geometries

2010 ◽  
Vol 133 (1) ◽  
Author(s):  
M. Eifel ◽  
V. Caspary ◽  
H. Hönen ◽  
P. Jeschke
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Base Flow ◽  
Internal Cooling ◽  
Turbine Rotor Blade ◽  
Flow Phenomena ◽  
Sst Turbulence Model

This paper presents the effects of major geometrical modifications to the interior of a convection cooled gas turbine rotor blade. The analysis of the flow is performed experimentally with flow visualization via paint injection into water, whereas the flow and the heat transfer are investigated numerically with ANSYS CFX, utilizing the SST turbulence model. Two sets of calculations are carried out: one under the same conditions as the experiments and another according to realistic hot gas conditions with conjugate heat transfer. The aim is to identify flow phenomena altering the heat transfer in the blade and to manipulate them in order to reduce the thermal load of the material. The operating point of the geometric base configuration is set to Re=50,000 at the inlet while for the modified geometries, the pressure ratio is held constant compared with the base. Flow structures and heat transfer conditions are evaluated and are linked to specific geometric features. Among several investigated configurations one could be identified that leads to a cooling effectiveness 15% larger compared with the base.

Comparison of Time-Resolved Turbine Rotor Blade Heat Transfer Measurements and Numerical Calculations

1992 ◽  
Vol 114 (4) ◽  
pp. 818-827 ◽  
Author(s):  
R. S. Abhari ◽  
G. R. Guenette ◽  
A. H. Epstein ◽  
M. B. Giles
Keyword(s):  
Heat Transfer ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Time History ◽  
Turbine Rotor ◽  
Numerical Calculations ◽  
Test Facility ◽  
Time Resolved ◽  
Turbine Rotor Blade ◽  
History Of

Time-resolved turbine rotor blade heat transfer data are compared with ab initio numerical calculations. The data were taken on a transonic, 4-to-1 pressure ratio, uncooled, single-stage turbine in a short-duration turbine test facility. The data consist of the time history of the heat transfer distribution about the rotor chord at midspan. The numerical calculation is a time accurate, two-dimensional, thin shear layer, multiblade row code known as UNSFLO. UNSFLO uses Ni’s Lax-Wendroff algorithm, conservative boundary conditions, and a time tilting algorithm to facilitate the calculation of the flow in multiple blade rows of arbitrary pitch ratio with relatively little computer time. The version used for this work had a simple algebraic Baldwin-Lomax turbulence model. The code is shown to do a good job of predicting the quantitative time history of the heat flux distribution. The wake/boundary layer and transonic interaction regions for suction and pressure surfaces are identified and the shortcomings of the current algebraic turbulence modeling in the code are discussed. The influence of hardware manufacturing tolerance on rotor heat transfer variation is discussed. A physical reasoning explaining the discrepancies between the unsteady measurement and the calculations for both the suction and pressure surfaces are given, which may be of use in improving future calculations and design procedures.

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.

Application of Artificial Neural Network for the Heat Transfer Investigation Around a High-Pressure Gas Turbine Rotor Blade

2011 ◽  
Author(s):  
Ibrahim Eryilmaz ◽  
Sinan Inanli ◽  
Baris Gumusel ◽  
Suha Toprak ◽  
Cengiz Camci
Keyword(s):  
Neural Network ◽  
Heat Transfer ◽  
Artificial Neural Network ◽  
High Pressure ◽  
Rotor Blade ◽  
Incidence Angle ◽  
Transfer Coefficients ◽  
Experimental Conditions ◽  
Heat Transfer Coefficients ◽  
Artificial Neural

This paper presents the preliminary results of using artificial neural networks in the prediction of gas side convective heat transfer coefficients on a high pressure turbine blade. The artificial neural network approach which has three hidden layers was developed and trained by nine inputs and it generates one output. Input and output data were taken from an experimental research program performed at the von Karman Institute for Fluid Dynamics by Camci and Arts [5,6] and Camci [7]. Inlet total pressure, inlet total temperature, inlet turbulence intensity, inlet and exit Mach numbers, blade wall temperature, incidence angle, specific location of measurement and suction/pressure side specification of the blade were used as input parameters and calculated heat transfer coefficient around a rotor blade used as output. After the network is trained with experimental data, heat transfer coefficients are interpolated for similar experimental conditions and compared with both experimental measurements and CFD solutions. CFD analysis was carried out to validate the algorithm and to determine heat transfer coefficients for a closely related test case. Good agreement was obtained between CFD results and neural network predictions.

scholarly journals Effect of Discrete Ribs on Heat Transfer and Friction Inside Narrow Rectangular Cross Section Cooling Passage of Gas Turbine Blade

2021 ◽  
Vol 10 (6) ◽  
pp. 192-209
Author(s):  
Karthik Krishnaswamy ◽  
◽  
Srikanth Salyan ◽  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Local Heat Transfer ◽  
Local Heat ◽  
Hydraulic Diameter ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Cooling Passage

The performance of a gas turbine during the service life can be enhanced by cooling the turbine blades efficiently. The objective of this study is to achieve high thermohydraulic performance (THP) inside a cooling passage of a turbine blade having aspect ratio (AR) 1:5 by using discrete W and V-shaped ribs. Hydraulic diameter (Dh) of the cooling passage is 50 mm. Ribs are positioned facing downstream with angle-of-attack (α) of 30° and 45° for discrete W-ribs and discerte V-ribs respectively. The rib profiles with rib height to hydraulic diameter ratio (e/Dh) or blockage ratio 0.06 and pitch (P) 36 mm are tested for Reynolds number (Re) range 30000-75000. Analysis reveals that, area averaged Nusselt numbers of the rib profiles are comparable, with maximum difference of 6% at Re 30000, which is within the limits of uncertainty. Variation of local heat transfer coefficients along the stream exhibited a saw tooth profile, with discrete W-ribs exhibiting higher variations. Along spanwise direction, discrete V-ribs showed larger variations. Maximum variation in local heat transfer coefficients is estimated to be 25%. For experimented Re range, friction loss for discrete W-ribs is higher than discrete-V ribs. Rib profiles exhibited superior heat transfer capabilities. The best Nu/Nuo achieved for discrete Vribs is 3.4 and discrete W-ribs is 3.6. In view of superior heat transfer capabilities, ribs can be deployed in cooling passages near the leading edge, where the temperatures are very high. The best THPo achieved is 3.2 for discrete V-ribs and 3 for discrete W-ribs at Re 30000. The ribs can also enhance the power-toweight ratio as they can produce high thermohydraulic performances for low blockage ratios.

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