Detailed Heat Transfer Coefficient Distributions on a Large-Scale Gas Turbine Blade Tip

2000 ◽  
Vol 123 (4) ◽  
pp. 803-809 ◽  
Author(s):  
Shuye Teng ◽  
Je-Chin Han ◽  
G. M. S. Azad
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Wind Tunnel ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Large Scale ◽  
Tip Clearance ◽  
Low Speed ◽  
Blade Tip ◽  
Low Speed Wind Tunnel

Measurements of detailed heat transfer coefficient distributions on a turbine blade tip were performed in a large-scale, low-speed wind tunnel facility. Tests were made on a five-blade linear cascade. The low-speed wind tunnel is designed to accommodate the 107.49 deg turn of the blade cascade. The mainstream Reynolds number based on cascade exit velocity was 5.3×105. Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. The central blade had a variable tip gap clearance. Measurements were made at three different tip gap clearances of about 1.1 percent, 2.1 percent, and 3 percent of the blade span. Static pressure distributions were measured in the blade mid-span and on the shroud surface. Detailed heat transfer coefficient distributions were measured on the blade tip surface using a transient liquid crystal technique. Results show that reduced tip clearance leads to reduced heat transfer coefficient over the blade tip surface. Results also show that reduced tip clearance tends to weaken the unsteady wake effect on blade tip heat transfer.

Experimental Evidence of Temperature Ratio Effect on Turbine Blade Tip Heat Transfer

2018 ◽  
Vol 140 (12) ◽  
Author(s):  
H. Jiang ◽  
Q. Zhang ◽  
L. He ◽  
S. Lu ◽  
L. Wang ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Boundary Conditions ◽  
Transfer Coefficient ◽  
Experimental Evidence ◽  
Turbine Blade ◽  
Tip Clearance ◽  
Temperature Ratio ◽  
Thermal Boundary Conditions ◽  
Thermal Measurements

Determination of a scalable Nusselt number (based on “adiabatic heat transfer coefficient”) has been the primary objective of the most existing heat transfer experimental studies. Based on the assumption that the wall thermal boundary conditions do not affect the flow field, the thermal measurements were mostly carried out at near adiabatic condition without matching the engine realistic wall-to-gas temperature ratio (TR). Recent numerical studies raised a question on the validity of this conventional practice in some applications, especially for turbine blade. Due to the relatively low thermal inertia of the over-tip-leakage (OTL) flow within the thin clearance, the fluids' transport properties vary greatly with different wall thermal boundary conditions and the two-way coupling between OTL aerodynamics and heat transfer cannot be neglected. The issue could become more severe when the gas turbine manufacturers are making effort to achieve much tighter clearance. However, there has been no experimental evidence to back up these numerical findings. In this study, transient thermal measurements were conducted in a high-temperature linear cascade rig for a range of tip clearance ratio (G/S) (0.3%, 0.4%, 0.6%, and 1%). Surface temperature history was captured by infrared thermography at a range of wall-to-gas TRs. Heat transfer coefficient (HTC) distributions were obtained based on a conventional data processing technique. The profound influence of tip surface thermal boundary condition on heat transfer and OTL flow was revealed by the first-of-its-kind experimental data obtained in the present experimental study.

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]

Unsteady Analysis on the Effects of Tip Clearance Height on Hot Streak Migration Across Rotor Blade Tip Clearance

2014 ◽  
Vol 136 (8) ◽  
Author(s):  
Zhaofang Liu ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Rotor Blade ◽  
Tip Clearance ◽  
Inlet Temperature ◽  
Leakage Flow ◽  
Blade Tip ◽  
Hot Streak ◽  
Blade Tip Clearance

This paper presents an investigation on the hot streak migration across rotor blade tip clearance in a high pressure gas turbine with different tip clearance heights. The blade geometry is taken from the first stage of GE-E3 turbine engine. Three tip clearances, 1.0%, 1.5%, and 2.5% of the blade span with a flat tip were investigated, respectively, and the uniform and nonuniform inlet temperature profiles were taken as the inlet boundary conditions. A new method for heat transfer coefficient calculation recommended by Maffulli and He has been adopted. By solving the unsteady compressible Reynolds-averaged Navier–Stokes equations, the time dependent solutions were obtained. The results indicate that the large tip clearance intensifies the leakage flow, increases the hot streak migration rate, and aggravates the heat transfer environment on the blade tip. However, the reverse secondary flow dominated by the relative motion of casing is insensitive to the change of tip clearance height. Attributed to the high-speed rotation of rotor blade and the low pressure difference between both sides of blade, a reverse leakage flow zone emerges over blade tip near trailing edge. Because it is possible for heat transfer coefficient distributions to be greatly different from heat flux distributions, it becomes of great concern to combine both of them in consideration of hot streak migration. To eliminate the effects of blade profile variation due to twist along the blade span on the aerothermal performance in tip clearance, the tested rotor (straight) blade and the original rotor (twisted) blade of GE-E3 first stage with the same tip profile are compared in this paper.

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.

Transonic Passage Turbine Blade Tip Clearance With Scalloped Shroud: Part III — Heat Transfer in Engine Configuration

2004 ◽  
Author(s):  
F. Casey Wilkins ◽  
Gregory M. Feldman ◽  
Wayne S. Strasser ◽  
James H. Leylek
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Large Scale ◽  
Numerical Study ◽  
Viscous Sublayer ◽  
Tip Clearance ◽  
Transfer Coefficients ◽  
Constant Wall Heat Flux ◽  
Blade Tip ◽  
Large Scale Land

This work presents a numerical study that was done to investigate the heat transfer characteristics of a transonic turbine blade with a scalloped shroud operating at realistic engine conditions typical of those found in a large scale, land-based gas turbine. The geometry under investigation was an infinite, linear cascade composed of the same blade and shroud design used in an experimental test rig by the research sponsor. This simulation was run for varying nominal tip clearances of 20, 80, and 5.08 mm. For each of these clearances, the simulation was run with and without the scrubbing effects of the outer casing, resulting in a total of six cases that could be used to determine the influence of tip clearance and relative casing motion on heat transfer. A high quality grid (ranging from approximately 10–12 million finite volumes depending on tip clearance) with y+ for first layer cells at or below 1.0 everywhere was used to resolve the flow down to the viscous sublayer. The “realizable” k-ε turbulence model was used for all cases. A constant wall heat flux was imposed on all the surrounding surfaces to obtain heat transfer data. Results produced include a full map of heat transfer coefficients for the suction and pressure surfaces of the blade as well as the tip, shroud, and outer casing for every case. Physical mechanisms responsible for the final heat transfer outcome for all six cases are documented.

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 casing purge flow on heat transfer and cooling performance of blade squealer tip for a gas turbine stage

2021 ◽  
pp. 095765092110403
Author(s):  
Shijie Jiang ◽  
Zhigang Li ◽  
Jun Li
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Three Dimensional ◽  
Tip Clearance ◽  
Local Cooling ◽  
Cold Air ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Blade Tip

The first stage of GE-E3 turbine is employed to investigate effect of casing purge flow upstream rotor blade tip. Three-dimensional Reynolds-averaged Navier-Stokes (RANS) equations and standard k-ω model are solved to obtain tip heat transfer simulations. The results reveal that: heat transfer coefficient of blade tip surface can be significantly reduced when casing purge flow is set. Tip averaged heat transfer coefficient of cases with and without swirly velocity casing purge flow decrease 3.5% and 3.4% compared with the case without casing purge flow. Compared with case which blowing ratio equals to 0.5, it can be found that averaged tip heat transfer coefficient of cases which blowing ratio equals to 1.0 and 1.5 decrease 2.3% and 1.8%, respectively. Setting blowing ratio as 1.0 can best cool tip surface without wasting cold air resources. Increasing rotating speed can induce cold air entering tip trailing region and improve local cooling effect. Flow structure inside the tip clearance are also revealed and discussed.

The Use of High Blockage Ribs to Enhance Heat Transfer Coefficient Distributions in a Model of an Integrally Cast Cooling Manifold

2006 ◽  
Author(s):  
Ioannis Ieronymidis ◽  
David R. H. Gillespie ◽  
Peter T. Ireland ◽  
Robert Kingston
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Large Scale ◽  
Thermal Contact ◽  
Maximum Level ◽  
Blade Cooling ◽  
Cooling Passage ◽  
The Heat Transfer Coefficient

In this paper detailed experimental measurements and computational predictions of heat transfer coefficient distributions in a large scale perspex model of a novel integrally cast blade cooling geometry are reported. In a gas turbine blade, the cooling passage investigated is integrally cast into the blade wall, providing good thermal contact with the outer surface of the turbine blade. Flow enters the racetrack passage through the root of the blade and exits to a central plenum through a series of nineteen transfer holes equally spaced in a staggered arrangement across the span of the blade. The Reynolds number changes continuously along the passage length because of the continuous ejection of fluid through a series of 19 transfer holes to the plenum. The smooth passage surface opposite is in closest proximity to the external surface, and this investigation has characterised the heat transfer coefficient on this surface at a range of engine representative inlet Reynolds numbers using a hybrid transient liquid crystal technique. The ability of three different rib configurations to enhance the heat transfer on this surface was also determined. Because the passage at engine scale is necessarily small, the rib height in all cases was 32.5% of the passage height. As the entire passage wetted surface is able to contribute to the blade cooling, and knowledge of the heat transfer coefficient distribution on the holed surfaces is crucial to prediction of blade life, a commercial CFD package, Fluent, was used to predict the heat transfer coefficient distributions on the holed surface, where there was no optical access during these tests. This also allowed investigation of additional rib configurations, and comparison of the pressure penalty associated with each design. The study showed that the turbulator configuration used allows the position and maximum level of heat transfer coefficient enhancement to be chosen by the engine designer. For the configurations tested heat transfer coefficient enhancement of up to 32% and 51% could be achieved on the holed surface and the ribbed surface respectively. For minimum additional pressure drop 45° ribs should be used.

The Effect of External Casing Impingement Cooling Manifold Standoff Distance on Casing Contraction for Thermal Control of Blade Tip Clearance

2017 ◽  
Vol 140 (2) ◽  
Author(s):  
Myeonggeun Choi ◽  
David R. H. Gillespie ◽  
Leo V. Lewis
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Thermal Control ◽  
Tip Clearance ◽  
Standoff Distance ◽  
Impingement Cooling ◽  
Assembly Tolerance ◽  
Blade Tip ◽  
Blade Tip Clearance

Thermal closure of the engine casing is widely used to minimize undesirable blade tip leakage flows thus improving jet engine performance. This may be achieved using an impingement cooling scheme on the external casing wall, provided by manifolds attached to the outside of the engine. The assembly tolerance of these components leads to variation in the standoff distance between the manifold and the casing, and its effects on casing contraction must be understood to allow build tolerance to be specified. For cooling arrangements with promising performance, the variation in closure with standoff distance of z/d = 1–6 were investigated through a mixture of extensive numerical modeling and experimental validation. A cooling manifold, typical of that adopted by several engine companies, incorporating three different arrays of short cooling holes (chosen from previous study by Choi et al. (2016, “The Relative Performance of External Casing Impingement Cooling Arrangements for Thermal Control of Blade Tip Clearance,” ASME J. Turbomach., 138(3), p. 031005.)) and thermal control dummy flanges were considered. Typical contractions of 0.5–2.2 mm are achieved from the 0.02–0.35 kg/s of the current casing cooling flows. The variation in heat transfer coefficient observed with standoff distance is much lower for the sparse array investigated compared to previous designs employing arrays typical of blade cooling configurations. The reason for this is explained through interrogation of the local flow field and resultant heat transfer coefficient. This implies that acceptable control of the circumferential uniformity of case cooling can be achieved with relatively large assembly tolerance of the manifold relative to the casing.

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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