Numerical Investigation of the Blade Tip and Overtip Casing Aerothermal Performance in a High Pressure Turbine Stage

2016 ◽  
Author(s):  
Kun Du ◽  
Zhigang Li ◽  
Jun Li
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Thermal Load ◽  
Rotor Blades ◽  
Turbine Stage ◽  
Cavity Depth ◽  
Blade Tip

In modern transonic gas turbine engines, the blade tip and overtip casing endures high thermal load, therefore these components are always subjected to thermal failures due to large unsteady heat flux. The unsteadiness is induced by the interaction of the rotor blades and periodic upstream wake of the vanes. The present study adopts a typical high pressure gas turbine stage (GE-E3 engine), and the computational domain consists of 1 high pressure stator vane and 2 rotor blades. The rotor blade in question has a squealer tip with a clearance gap about 1% of the blade height. This study focuses on the physics of the heat transfer characteristic of the blade tip and overtip casing regions. The present simulations were conducted using three-dimensional unsteady Reynolds-averaged Navier-Stokes (URANS) commercial code at the real engine conditions ( Mexit = 1.07, n = 8450rpm ). The standard k–ω turbulence model was utilized to model the turbulence. The accuracy of CFD predictions has been validated by comparison with the experimental data. The steady, unsteady and time-averaged results on the blade tip and overtip casing have been observed and discussed. Results indicate that the depth of the cavity has great influence on the blade tip and overtip casing. The averaged heat transfer coefficient on the blade tip is reduced with the increase of the cavity depth, however, the thermal load on the blade tip presents a contrary tendency. Moreover, the largest unsteadiness was observed for the case with D = 3.0 among the cases investigated, especially near the suction side squealer. In addition, the variation of the cavity depth has little effect on the heat transfer coefficient and thermal load on the overtip casing.

scholarly journals Heat Transfer and Flow on the Squealer Tip of a Gas Turbine Blade

2000 ◽  
Author(s):  
Gm S. Azad ◽  
Je-Chin Han ◽  
Robert J. Boyle
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Rotor Blade ◽  
Static Pressure ◽  
Cavity Surface ◽  
Edge Region ◽  
Blade Tip ◽  
The Heat Transfer Coefficient

Experimental investigations are performed to measure the detailed heat transfer coefficient and static pressure distributions on the squealer tip of a gas turbine blade in a five-bladed stationary linear cascade. The blade is a 2-dimensional model of a modern first stage gas turbine rotor blade with a blade tip profile of a GE-E3 aircraft gas turbine engine rotor blade. A squealer (recessed) tip with a 3.77% recess is considered here. The data on the squealer tip are also compared with a flat tip case. All measurements are made at three different tip gap clearances of about 1%, 1.5%, and 2.5% of the blade span. Two different turbulence intensities of 6.1% and 9.7% at the cascade inlet are also considered for heat transfer measurements. Static pressure measurements are made in the mid-span and near-tip regions, as well as on the shroud surface opposite to the blade tip surface. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and an exit Reynolds number based on the axial chord of 1.1×106. A transient liquid crystal technique is used to measure the heat transfer coefficients. Results show that the heat transfer coefficient on the cavity surface and rim increases with an increase in tip clearance. The heat transfer coefficient on the rim is higher than the cavity surface. The cavity surface has a higher heat transfer coefficient near the leading edge region than the trailing edge region. The heat transfer coefficient on the pressure side rim and trailing edge region is higher at a higher turbulence intensity level of 9.7% over 6.1% case. However, no significant difference in local heat transfer coefficient is observed inside the cavity and the suction side rim for the two turbulence intensities. The squealer tip blade provides a lower overall heat transfer coefficient when compared to the flat tip blade.

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]

Heat Transfer Coefficient and Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2002 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Rotor Blade ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

The detailed distributions of heat transfer coefficient and film cooling effectiveness on a gas turbine blade tip were measured using a hue detection based transient liquid crystal technique. 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 the GE-E3 aircraft gas turbine engine rotor blade. 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 97.7°. The overall pressure ratio was 1.32 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 blade model was equipped with a single row of film cooling holes at both the tip portion along the camber line and near the tip region of the pressure-side. All measurements were made at the three different tip gap clearances of 1%, 1.5%, and 2.5% of blade span and the three blowing ratios of 0.5, 1.0, and 2.0. Results showed that, in general, heat transfer coefficient and film effectiveness increased with increasing tip gap clearance. As blowing ratio increased, heat transfer coefficient decreased, while film effectiveness increased. Results also showed that adding pressure-side coolant injection would further decrease blade tip heat transfer coefficient but increase film effectiveness.

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.

Experimental Survey on Heat Transfer in a Trailing Edge Cooling System: Effects of Rotation in Internal Cooling Ducts

2012 ◽  
Author(s):  
L. Bonanni ◽  
C. Carcasci ◽  
B. Facchini ◽  
L. Tarchi
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Cooling System ◽  
Internal Cooling ◽  
Test Rig ◽  
Trailing Edge ◽  
Tip Mass

The high thermal loads, the heavy structural stresses and the small thickness required for aerodynamic performances make the trailing edge cooling (TE) cooling of high pressure gas turbine blades a critical challenge. The presented paper point out an experimental study focusing the aerothermal performance of a TE internal cooling system of a high pressure gas turbine blade, evaluated under stationary and rotating conditions. The investigated geometry consists of a 30:1 scaled model reproducing the typical wedge shaped discharge duct with one row of enlarged pedestals. The airflow pattern inside the device simulates a highly loaded rotor blade cooling scheme with a 90° turning flow from the radial hub inlet to the tangential TE outlet. Two different tip configurations were tested, the first one with a completely closed section, the second one with 5 holes on the tip outlet surfaces discharging at ambient pressure. To investigate the rotation effects on the trailing edge cooling system performance, a rotating test rig was purposely developed and manufactured. The test rig is composed by a rotating arm that holds the PMMA TE model and the instrumentation. A thin Inconel heating foil and wide band Thermo-chromic Liquid Crystals are used to perform steady state heat transfer measurements. A rotary joint ensures the pneumatic connection between the blower and the rotating apparatus, moreover several slip rings are used for both instrumentation power supply and thermocouple connection. Heat transfer coefficient measurements were made with fixed Reynolds number close to 20k in the hub inlet section and with variable rotating speed in order to set the Rotation number from 0 (non rotational test) up to 0.3. Six different configurations were tested: two different tip mass flow rates (the first one with a completely closed tip, the second one with the 12.5% of the inlet flow discharged from the tip) and three different surface conditions: the first one consists in the flat plate case and the others in two ribbed cases, with different angular orientation (60° and −60° respect to the radial direction). Results are reported in terms of detailed heat transfer coefficient 2D maps on the suction side surface as well as span-wise profiles inside the pedestal ducts. The reported work has been supported by the Italian Ministry of Education, University and Research (MIUR).

Heat Transfer Coefficients on the Squealer Tip and Near Tip Regions of a Gas Turbine Blade With Single or Double Squealer

2003 ◽  
Author(s):  
Jae Su Kwak ◽  
Jaeyong Ahn ◽  
Je-Chin Han ◽  
C. Pang Lee ◽  
Robert Boyle ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Blade Tip

Detailed heat transfer coefficient distributions on a gas turbine squealer tip 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. Squealer rims were located along (a) the camber line, (b) the pressure side, (c) the suction side, (d) the pressure and suction sides, (e) the camber line and the pressure side, and (f) the camber line and the suction side, respectively. Tests were performed on a five-bladed linear cascade with a blow down facility. The Reynolds number based on the cascade exit velocity and the axial chord length of a blade was 1.1×106 and the overall pressure ratio was 1.2. Heat transfer measurements were taken at the three tip gap clearances of 1.0%, 1.5% and 2.5% of blade span. Results show that the heat transfer coefficients on the blade tip and the shroud were significantly reduced by using a squealer tip blade. Results also showed that a different squealer geometry arrangement changed the leakage flow path and resulted in different heat transfer coefficient distributions. The suction side squealer tip provided the lowest heat transfer coefficient on the blade tip and near tip regions compared to the other squealer geometry arrangements.

Unsteady Passing Wake Effects on Turbine Blade Tip Aerodynamic and Aerothermal Performance With Film Cooling

2019 ◽  
Author(s):  
Bo Zhang ◽  
Xiaoqing Qiang ◽  
Jinfang Teng ◽  
Shaopeng Lu
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Guide Vane ◽  
Rotor Blades ◽  
Leakage Flow ◽  
Wake Effects ◽  
Blade Tip ◽  
Film Cooling Holes

Abstract In this paper, upstream unsteady passing wake effects on the rotor blade tip with film cooling have been numerically examined. The geometry and flow conditions of the first stage of GE-E3 high-pressure turbine have been used to obtain the unsteady three-dimensional blade tip flow and heat transfer characteristics. The first stage of GE-E3 high-pressure turbine has 46 guide vanes and 76 rotor blades. In the process of calculation, compromise the computational resources and accuracy to simplify the number of the guide vane and rotor blade to 38:76. Namely, each computational domain comprises of one guide vane and two rotor blades. The computational boundary conditions are consistent with the GE-E3 annular cascade test conditions. In the case of 1% blade span tip clearance, through comparing the steady results, time-averaged and time-resolved unsteady results, the investigation of the unsteady passing wake effects on flat tip aerodynamic and aerothermal performance without film cooling holes (NC) and three cases with film cooling holes are conducted. To be specific, near the leading edge (C123), at the central area (C456) and near the trailing edge (C789). This paper emphasizes the variation of leakage flow and heat transfer coefficient at different unsteady instants. The results show that the time-averaged leakage flow is pretty similar to the steady results, but the increment of the leakage flow can rise to more than 8% at the maximum envelope. Moreover, the heat transfer coefficient discrepancy between steady results and time-averaged results is almost below 4%, but the dramatic growth of the instantaneous heat transfer coefficient along the pressure side is in excess of 20% due to the shift of the pressure spot.

The Effects of Blade Passing on the Heat Transfer Coefficient of the Overtip Casing in a Transonic Turbine Stage

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
Steven J. Thorpe ◽  
Roger W. Ainsworth
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Cooling System ◽  
Turbine Stage ◽  
Transfer Coefficients ◽  
Time Resolved ◽  
Turbine Engine ◽  
Transonic Turbine

In a modern gas turbine engine, the outer casing (shroud) of the shroudless high-pressure turbine is exposed to a combination of high flow temperatures and heat transfer coefficients. The casing is consequently subjected to high levels of convective heat transfer, a situation that is complicated by flow unsteadiness caused by periodic blade-passing events. In order to arrive at an overtip casing design that has an acceptable service life, it is essential for manufacturers to have appropriate predictive methods and cooling system configurations. It is known that both the flow temperature and boundary layer conductance on the casing wall vary during the blade-passing cycle. The current article reports the measurement of spatially and temporally resolved heat transfer coefficient (h) on the overtip casing wall of a fully scaled transonic turbine stage experiment. The results indicate that h is a maximum when a blade tip is immediately above the point in question, while the lower values of h are observed when the point is exposed to the rotor passage flow. Time-resolved measurements of static pressure are used to reveal the unsteady aerodynamic situation adjacent to the overtip casing wall. The data obtained from this fully scaled transonic turbine stage experiment are compared to previously published heat transfer data obtained in low-Mach number cascade-style tests of similar high-pressure blade geometries.

Turbine Blade Tip Cooling Modeling and Optimization

2009 ◽  
Author(s):  
Gregory Vogel ◽  
Anmol Agrawal ◽  
Praneetha Nannapaneni
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Metal Temperature ◽  
Cfd Analysis ◽  
Tip Leakage ◽  
Cooling Flow ◽  
Blade Tip

The turbine blade tip is considered as one of the most critical areas of gas turbine engines. The tip region often lacks durability and is challenging to cool. To achieve successful blade tip cooling designs, ALSTOM engineers are performing state of the art aero thermal analyses of blade tip cooling configurations. This paper describes the approach used for this analysis and draws conclusion for blade tip cooling optimization. Numerical simulations of flow and heat transfer are presented for a modern industrial gas turbine blade with a film cooled tip. The blade tip metal temperature distribution is analyzed for three different blade tip clearances with a detailed CFD analysis around the blade tip performed. The CFD analysis provides flow streamlines through the blade tip as well as a total blade tip leakage flow. Rough streamlines estimates are then used to define a set of control volumes for which dedicated cooling flow mixing is considered. The total mass flowing through all volumes corresponds to the CFD blade tip leakage. For each control volume corresponds a specific Reynolds number that is used to define a corresponding heat transfer coefficient. The latter is obtained from experimental Nusselt number correlations for the different regions of a blade squealer tip (crown, fillet and cavity). Application of the obtained heat transfer coefficient and mixing temperature boundary conditions on a 3D blade tip finite element model, together with an internal cooling flow network associated to the 3D model allows calculating the blade tip metal temperature. Results for two different tip clearances relative to nominal blade tip gap are presented and discussed. Comparison with experimental data such as thermal paint test and metallurgical data are given, showing good agreement with the blade tip cooling modeling introduced in this paper. Cooling performance of the blade tip is discussed based on the modeling approach proposed in this paper. The latter allows drawing conclusions for blade tip cooling optimization.

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