Heat Transfer and Effectiveness on the Film Cooled Tip and Inner Rim Surfaces of a Turbine Blade

2010 ◽  
Author(s):  
Jun Su Park ◽  
Dong Hyun Lee ◽  
Hyung Hee Cho ◽  
Dong-Ho Rhee ◽  
Shin-Hyung Kang
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
High Heat ◽  
Chord Length ◽  
Tip Clearance ◽  
Film Cooling Effectiveness ◽  
High Heat Transfer ◽  
Film Cooling Holes

Detailed heat/mass transfer coefficients and film-cooling effectiveness were measured on the tip and inner rim surfaces of a rotor blade with a squealer rim. The blade was a two-dimensional version of a modern first-stage gas turbine rotor blade with a squealer rim. The experimental apparatus was equipped with a linear cascade of three blades, the axial chord length (Cx) of which was 237 mm with a turning angle of 126°. The mainstream Reynolds number based on the axial chord was 1.5×105. The turbulence intensity level at the cascade inlet was approximately 12%. Measurements were made at three different rim heights (H) of about 3%, 6%, and 9% of the axial chord length. The tip clearance (C) ranges were 1–3% of the axial chord length. Also, three different types of blade tip surfaces were equipped with a single row of film-cooling holes along the camber line, near the pressure and the suction side rim. In particular, a coolant was injected at an incline of 45° from near the suction side film cooling holes. The film cooling experiments were done with a fixed tip clearance and rim height at 1% and 6% of the axial chord length. The blowing rate was fixed at 1.5. High heat transfer rates were observed near the leading edge on the tip surface in some cases, due to the reattachment of tip leakage flow. The peak values moved toward the suction-side edge, and the magnitude and area of high heat transfer increased near the leading edge as the tip clearance increased. The heat transfer decreased on the tip surface with increases in the rim height. In the film-cooling cases, the high heat transfer and film-cooling effectiveness region appeared near the film-cooling holes.

scholarly journals Heat Transfer on a Film-Cooled Rotating Blade

1999 ◽  
Author(s):  
Vijay K. Garg
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Three Dimensional ◽  
Leading Edge ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

A multi-block, three-dimensional Navier-Stokes code has been used to compute heat transfer coefficient on the blade, hub and shroud for a rotating high-pressure turbine blade with 172 film-cooling holes in eight rows. Film cooling effectiveness is also computed on the adiabatic blade. Wilcox’s k-ω model is used for modeling the turbulence. Of the eight rows of holes, three are staggered on the shower-head with compound-angled holes. With so many holes on the blade it was somewhat of a challenge to get a good quality grid on and around the blade and in the tip clearance region. The final multi-block grid consists of 4784 elementary blocks which were merged into 276 super blocks. The viscous grid has over 2.2 million cells. Each hole exit, in its true oval shape, has 80 cells within it so that coolant velocity, temperature, k and ω distributions can be specified at these hole exits. It is found that for the given parameters, heat transfer coefficient on the cooled, isothermal blade is highest in the leading edge region and in the tip region. Also, the effectiveness over the cooled, adiabatic blade is the lowest in these regions. Results for an uncooled blade are also shown, providing a direct comparison with those for the cooled blade. Also, the heat transfer coefficient is much higher on the shroud as compared to that on the hub for both the cooled and the uncooled cases.

Predictions of Cooling From Dirt Purge Holes Along the Tip of a Turbine Blade

2003 ◽  
Author(s):  
E. M. Hohlfeld ◽  
J. R. Christophel ◽  
E. L. Couch ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Suction Side ◽  
High Heat ◽  
Blowing Ratio ◽  
High Heat Transfer ◽  
Blade Tip ◽  
Flow Through ◽  
Film Cooling Holes

The clearance gap between the tip of a turbine blade and its associated shroud provides a flow path for leakage from the pressure side of the blade to the suction side. The tip region is one area that experiences high heat transfer and, as such, can be the determining factor for blade life. One method for reducing blade tip heat transfer is to use cooler fluid from the compressor, that exits from relatively large dirt purge holes placed in the tip, for cooling purposes. Dirt purge holes are typically manufactured in the blade tip to extract dirt from the coolant flow through centrifugal forces such that these dirt particles do not block smaller diameter film-cooling holes. This paper discusses the results of numerous computational simulations of cooling injection from dirt purge holes along the tip of a turbine blade. Some comparisons are also made to experimental results in which a properly scaled-up blade geometry (12X) was used to form a two-passage linear cascade. Computational results indicate that the cooling achieved through the dirt purge injection from the blade tip is dependent on the gap size as well as the blowing ratio. For a small tip gap (0.54% of the span) the flow exiting the dirt purge holes act as a blockage for the leakage flow across the gap. As the blowing ratio is increased for a large tip gap (1.63% of the span), the tip cooling increases only slightly while the cooling to the shroud increases significantly.

Turbine Blade Tip Film Cooling With Blade Rotation: Part I — Tip and Pressure Side Coolant Injection

2015 ◽  
Author(s):  
Onieluan Tamunobere ◽  
Sumanta Acharya
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
High Heat ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
High Heat Transfer ◽  
Blade Tip

This is the first in a two-part series of an experimental film cooling study conducted on the tip of a turbine blade with a blade rotation speed of 1200 RPM. In this part of the study, the coolant is injected from the blade tip and pressure side (PS) holes, and the effect of the blowing ratio on the heat transfer coefficient and film cooling effectiveness of the blade tip is investigated. The blade has a tip clearance of 1.7% of the blade span and consists of a cut back squealer rim, two cylindrical tip holes and six shaped pressure side holes. The stator-rotor-stator test section is housed in a closed loop wind tunnel that allows for the performance of transient heat transfer tests. Measurements of the heat transfer coefficient and film cooling effectiveness are done on the blade tip using liquid crystal thermography. These measurements are reported for the no coolant case and for blowing ratios of 1.0, 1.5, 2.0, 3.0 and 4.0. The heat transfer results for the no coolant injection show a region of high heat transfer on the blade tip near the blade leading edge region as the incident flow impinges on that region. This region of high heat transfer extends and stretches on the tip as more coolant is introduced through the tip holes at higher blowing ratios. The cooling results show that increasing the blowing ratio increases the film cooling effectiveness. The tip film cooling profile is such that the tip coolant is pushed towards the blade suction side thereby providing better coverage in that region. The shift in coolant flow profile towards the blade suction side as opposed to the pressure side in stationary studies can primarily be attributed to the effects of the blade relative motion.

Numerical Prediction of Film Cooling and Heat Transfer on the Leading Edge of a Rotating Blade in a 1-1/2 Turbine Stage

2004 ◽  
Author(s):  
Huitao Yang ◽  
Hamn-Ching Chen ◽  
Je-Chin Han ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Gas Turbines ◽  
Turbine Blades ◽  
Leading Edge ◽  
High Heat ◽  
Turbine Stage ◽  
Film Cooling Effectiveness ◽  
Rotating Blade ◽  
High Heat Transfer

In modern gas turbines, the blade leading edge region is one area that experiences high heat transfer due to the stagnation flow. Many cooling techniques have been applied to blades, so they can withstand these high heat loads; one of the common methods in cooling turbine blades is to apply film cooling. In the present study, numerical simulations were performed to predict the film cooling effectiveness and heat transfer coefficient on the leading edge of a rotating blade in a 1-1/2 turbine stage using a Reynolds stress turbulence model together with a non-equilibrium wall function. In addition, the unsteady characteristics of the film cooling and heat transfer at different time phases during a passing period were also investigated.

Turbine Blade Tip Film Cooling With Blade Rotation—Part I: Tip and Pressure Side Coolant Injection

2016 ◽  
Vol 138 (9) ◽  
Author(s):  
Onieluan Tamunobere ◽  
Sumanta Acharya
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
High Heat ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
High Heat Transfer ◽  
Blade Tip ◽  
The Heat Transfer Coefficient

An experimental study of film cooling is conducted on the tip of a turbine blade with a blade rotation speed of 1200 rpm. The coolant is injected from the blade tip and pressure side (PS) holes, and the effect of the blowing ratio on the heat transfer coefficient and film cooling effectiveness of the blade tip is investigated. The blade has a tip clearance of 1.7% of the blade span and consists of a cut back squealer rim, two cylindrical tip holes, and six shaped PS holes. The stator–rotor–stator test section is housed in a closed loop wind tunnel that allows for the performance of transient heat transfer tests. Measurements of the heat transfer coefficient and film cooling effectiveness are done on the blade tip using liquid crystal thermography. These measurements are reported for the no coolant case and for blowing ratios of 1.0, 1.5, 2.0, 3.0, and 4.0. The heat transfer result for the no coolant injection shows a region of high heat transfer on the blade tip near the blade leading edge region as the incident flow impinges on that region. This region of high heat transfer extends and stretches on the tip as more coolant is introduced through the tip holes at higher blowing ratios. The cooling results show that increasing the blowing ratio increases the film cooling effectiveness. The cooling effectiveness signatures indicate that the tip coolant is pushed toward the blade suction side thereby providing better coverage in that region. This shift in coolant flow toward the blade suction side, as opposed to the PS in stationary studies, can primarily be attributed to the effects of the blade relative motion.

Influence of tip clearance and cavity depth on heat transfer in a cutback squealer tip

2020 ◽  
pp. 095441002096469
Author(s):  
Weijie Wang ◽  
Shaopeng Lu ◽  
Hongmei Jiang ◽  
Qiusheng Deng ◽  
Jinfang Teng ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Suction Side ◽  
High Heat ◽  
Tip Clearance ◽  
Tip Leakage Flow ◽  
Trailing Edge ◽  
Cavity Depth ◽  
High Heat Transfer ◽  
Blade Tip ◽  
High Heat Transfer Coefficient

Numerical simulations are conducted to present the aerothermal performance of a turbine blade tip with cutback squealer rim. Two different tip clearance heights (0.5%, 1.0% of the blade span) and three different cavity depths (2.0%, 3.0%, and 6.0% of the blade span) are investigated. The results show that a high heat transfer coefficient (HTC) strip on the cavity floor appears near the suction side. It extends with the increase of tip clearance height and moves towards the suction side with the increase of cavity depth. The cutback region near the trailing edge has a high HTC value due to the flush of over-tip leakage flow. High HTC region shrinks to the trailing edge with the increase of cavity depth since there is more accumulated flow in the cavity for larger cavity depth. For small tip clearance cases, high HTC distribution appears on the pressure side rim. However, high HTC distribution is observed on suction side rim for large tip clearance height. This is mainly caused by the flow separation and reattachment on the squealer rims.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

Platform Film Cooling Investigation on an HP Nozzle Vane Cascade With Discrete Shaped Holes and Slot Film Cooling

2020 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
L. Abba ◽  
L. Pestelli
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Heat Transfer Performance ◽  
Leading Edge ◽  
Suction Side ◽  
Secondary Flows ◽  
Intensity Level ◽  
Transfer Performance ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

Abstract The present paper reports on an experimental investigation on the aerodynamic and heat transfer performance of different platform cooling schemes: two based on cylindrical and shaped holes and one featuring a slot located upstream of the leading edge plane simulating the combustor to stator interface gap. Tests were run on a 6-vane cascade operated at an isentropic cascade exit Mach number of 0.4 and a significant inlet turbulence intensity level of about 9%. The cooling schemes were first tested to quantify their impact on secondary flows and related losses for variable injection conditions. Heat transfer performance was then assessed through adiabatic film cooling effectiveness and heat transfer coefficient measurements. The Net Heat Flux Reduction parameter was then computed to critically assess the cooling schemes. When compared with the cylindrical hole scheme, shaped holes outperform for all tested injection rates, while the slot alone is able to thermally protect only the front of the passage. Discrete holes are required to cool the platform region along the whole pressure side and the suction side leading edge region.

Bump and Trench Modifications to Film-Cooling Holes at the Vane-Endwall Junction

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Junction Region ◽  
High Heat Transfer ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

The endwall of a first-stage vane experiences high heat transfer and low adiabatic effectiveness levels because of high turbine operating temperatures and formation of leading edge vortices. These vortices lift the coolant off the endwall and pull the hot mainstream gases toward it. The region of focus for this study is the vane-endwall junction region near the stagnation location where cooling is very difficult. Two different film-cooling hole modifications, namely, trenches and bumps, were evaluated to improve the cooling in the leading edge region. This study uses a large-scale turbine vane cascade with a single row of axial film-cooling holes at the leading edge of the vane endwall. Individual hole trenches and row trenches were placed along the complete row of film-cooling holes. Two-dimensional semi-elliptically shaped bumps were also evaluated by placing the bumps upstream and downstream of the film-cooling row. Tests were carried out for different trench depths and bump heights under varying blowing ratios. The results indicated that a row trench placed along the row of film-cooling holes showed a greater enhancement in adiabatic effectiveness levels when compared to individual hole trenches and bumps. All geometries considered produced an overall improvement to adiabatic effectiveness levels.

Bump and Trench Modifications to Film-Cooling Holes at the Vane Endwall Junction

2007 ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Junction Region ◽  
High Heat Transfer ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

The endwall of a first stage vane experiences high heat transfer and low adiabatic effectiveness levels because of high turbine operating temperatures and formation of leading edge vortices. These vortices lift the coolant off the endwall and pull the hot mainstream gases towards it. The region of focus for this study is the vane-endwall junction region near the stagnation location where cooling is very difficult. Two different film-cooling hole modifications, namely trenches and bumps, were evaluated to improve the cooling in the leading edge region. This study uses a large-scale turbine vane cascade with a single row of axial film-cooling holes at the leading edge of the vane endwall. Individual hole trenches and row trenches were placed along the complete row of film-cooling holes. Two-dimensional semi-elliptically shaped bumps were also evaluated by placing the bumps upstream and downstream of the film-cooling row. Tests were carried out for different trench depths and bump heights under varying blowing ratios. The results indicated that a row trench placed along the row of film-cooling holes showed a greater enhancement in adiabatic effectiveness levels when compared to individual hole trenches and bumps. All geometries considered produced an overall improvement to adiabatic effectiveness levels.

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