scholarly journals Infrared Low-Temperature Turbine Vane Rough Surface Heat Transfer Measurements

2000 ◽  
Vol 123 (1) ◽  
pp. 168-177 ◽  
Author(s):  
R. J. Boyle ◽  
C. M. Spuckler ◽  
B. L. Lucci ◽  
W. P. Camperchioli
Keyword(s):  
Heat Transfer ◽  
Infrared Camera ◽  
Leading Edge ◽  
Suction Side ◽  
Surface Heat ◽  
Test Facility ◽  
Edge Region ◽  
Linear Cascade ◽  
Test Conditions ◽  
Turbine Vane

Turbine vane heat transfer distributions obtained using an infrared camera technique are described. Infrared thermography was used because noncontact surface temperature measurements were desired. Surface temperatures were 80°C or less. Tests were conducted in a three-vane linear cascade, with inlet pressures between 0.14 and 1.02 atm, and exit Mach numbers of 0.3, 0.7, and 0.9, for turbulence intensities of approximately 1 and 10 percent. Measurements were taken on the vane suction side, and on the pressure side leading edge region. The designs for both the vane and test facility are discussed. The approach used to account for conduction within the vane is described. Midspan heat transfer distributions are given for the range of test conditions.

scholarly journals Infrared Low Temperature Turbine Vane Rough Surface Heat Transfer Measurements

2000 ◽  
Author(s):  
R. J. Boyle ◽  
C. M. Spuckler ◽  
B. L. Lucci ◽  
W. P. Camperchioli
Keyword(s):  
Heat Transfer ◽  
Infrared Camera ◽  
Leading Edge ◽  
Suction Side ◽  
Surface Heat ◽  
Test Facility ◽  
Edge Region ◽  
Linear Cascade ◽  
Test Conditions ◽  
Turbine Vane

Turbine vane heat transfer distributions obtained using an infrared camera technique are described. Infrared thermography was used because non-contact surface temperature measurements were desired. Surface temperatures were 80°C or less. Tests were conducted in a three vane linear cascade, with inlet pressures between 0.14 and 1.02atm., and exit Mach numbers of 0.3, 0.7, and 0.9, for turbulence intensities of approximately 1 and 10%. Measurements were taken on the vane suction side, and on the pressure side leading edge region. The designs for both the vane and test facility are discussed. The approach used to account for conduction within the vane is described. Midspan heat transfer distributions are given for the range of test conditions.

Evaluation of Superposition Predictions for Showerhead Film Cooling on a Vane

2014 ◽  
Author(s):  
Joshua B. Anderson ◽  
James R. Winka ◽  
David G. Bogard ◽  
Michael E. Crawford
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Edge Region ◽  
Current Investigation ◽  
Engine Design ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

The leading edge of a turbine vane is subject to some of the highest temperature loading within an engine, and an accurate understanding of leading edge film coolant behavior is essential for modern engine design. Although there have been many investigations of the adiabatic effectiveness for showerhead film cooling of a vane leading edge region, there have been no previous studies in which individual rows of the showerhead were tested with the explicit intent of validating superposition models. For the current investigation, a series of adiabatic effectiveness experiments were performed with a five-row and three-row showerhead. The experiments were repeated separately with each individual row of holes active. This allowed evaluation of superposition methods on both the suction side of the vane, which was moderately convex, and the pressure side of the vane, which was mildly concave. Superposition was found to accurately predict performance on the suction side of the vane at lower momentum flux ratios, but not at higher momentum flux ratios. On the pressure side of the vane the superposition predictions were consistently lower than measured values, with significant errors occurring at the higher momentum flux ratios. Reasons for the under-prediction by superposition analysis are presented.

A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall: Part I — Introduction and Aerodynamic Measurements

2015 ◽  
Author(s):  
Franz Puetz ◽  
Johannes Kneer ◽  
Achmed Schulz ◽  
Hans-Joerg Bauer
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Film Cooling ◽  
Gas Turbines ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Test Facility ◽  
Endwall Contouring

An increased demand for lower emission of stationary gas turbines as well as civil aircraft engines has led to new, low emission combustor designs with less liner cooling and a flattened temperature profile at the outlet. As a consequence, the heat load on the endwall of the first nozzle guide vane is increased. The secondary flow field dominates the endwall heat transfer, which also contributes to aerodynamic losses. A promising approach to reduce these losses is non-axisymmetric endwall contouring. The effects of non-axisymmetric endwall contouring on heat transfer and film cooling are yet to be investigated. Therefore, a new cascade test rig has been set up in order to investigate endwall heat transfer and film cooling on both a flat and a non-axisymmetric contoured endwall. Aerodynamic measurements that have been made prior to the upcoming heat transfer investigation are shown. Periodicity and detailed vane Mach number distributions ranging from 0 to 50% span together with the static pressure distribution on the endwall give detailed information about the aerodynamic behavior and influence of the endwall contouring. The aerodynamic study is backed by an oil paint study, which reveals qualitative information on the effect of the contouring on the endwall flow field. Results show that the contouring has a pronounced effect on vane and endwall pressure distribution and on the endwall flow field. The local increase and decrease of velocity and the reduced blade loading towards the endwall is the expected behavior of the 3d contouring. So are the results of the oil paint visualization, which show a strong change of flow field in the leading edge region as well as that the contouring delays the horse shoe vortex hitting the suction side.

Film Cooling Jet Injection Effect in Heat Transfer Coefficient Augmentation for the Pressure Side Cooling of Turbine Vane

2014 ◽  
Author(s):  
Hossein Nadali Najafabadi ◽  
Matts Karlsson ◽  
Mats Kinell ◽  
Esa Utriainen
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Time Resolved ◽  
Turbine Vane

Improving film cooling performance of turbine vanes and blades is often achieved through application of multiple arrays of cooling holes on the suction side, the showerhead region and the pressure side. This study investigates the pressure side cooling under the influence of single and multiple rows of cooling in the presence of a showerhead from a heat transfer coefficient augmentation perspective. Experiments are conducted on a prototype turbine vane working at engine representative conditions. Transient IR thermography is used to measure time-resolved surface temperature and the semi-infinite method is utilized to calculate the heat transfer coefficient on a low conductive material. Investigations are performed for cylindrical and fan-shaped holes covering blowing ratio 0.6 and 1.8 at density ratio of about unity. The freestream turbulence is approximately 5% close to the leading edge. The resulting heat transfer coefficient enhancement, the ratio of HTC with to that without film cooling, from different case scenarios have been compared to showerhead cooling only. Findings of the study highlight the importance of showerhead cooling to be used with additional row of cooling on the pressure side in order to reduce heat transfer coefficient enhancement. In addition, it is shown that extra rows of cooling will not significantly influence heat transfer augmentation, regardless of the cooling hole shape.

Turbine blade heat transfer calculations using two-equation turbulence models

1997 ◽  
Vol 211 (3) ◽  
pp. 253-262 ◽  
Author(s):  
J Larsson
Keyword(s):  
Heat Transfer ◽  
Turbulence Models ◽  
Leading Edge ◽  
Suction Side ◽  
External Heat ◽  
Navier Stokes ◽  
Edge Region ◽  
Turbine Cascades ◽  
Good Agreement

A full Navier-Stokes solver is used to calculate external heat transfer in two linear two-dimensional turbine cascades, one subsonic and one transonic. Heat transfer results obtained with two low-Reynolds k-∊ models (Chien and Launder-Sharma) and two k-ω models (Wilcox standard and transition) are compared with measurements. Good agreement is found in some regions, but the suction side transition and the leading edge are not predicted correctly. Problems with turbulence levels that are too high in the leading edge region are investigated. The numerical quality of the results is investigated and a few general guidelines about the numerics are given. Grid and scheme independence is also demonstrated.

Evaluation of Superposition Predictions for Showerhead Film Cooling on a Vane

2015 ◽  
Vol 137 (4) ◽  
Author(s):  
Joshua B. Anderson ◽  
James R. Winka ◽  
David G. Bogard ◽  
Michael E. Crawford
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Edge Region ◽  
Current Investigation ◽  
Engine Design ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

The leading edge of a turbine vane is subject to some of the highest temperature loading within an engine, and an accurate understanding of leading edge film coolant behavior is essential for modern engine design. Although there have been many investigations of the adiabatic effectiveness for showerhead film cooling of a vane leading edge region, there have been no previous studies in which individual rows of the showerhead were tested with the explicit intent of validating superposition models. For the current investigation, a series of adiabatic effectiveness experiments were performed with a five-row and three-row showerhead. The experiments were repeated separately with each individual row of holes active. This allowed evaluation of superposition methods on both the suction side of the vane, which was moderately convex, and the pressure side of the vane, which was mildly concave. Superposition was found to accurately predict performance on the suction side of the vane at lower momentum flux ratios, but not at higher momentum flux ratios. On the pressure side of the vane, the superposition predictions were consistently lower than measured values, with significant errors occurring at the higher momentum flux ratios. Reasons for the underprediction by superposition analysis are presented.

On the heat transfer characteristics of a Lamilloy cooling structure with curvatures with different pin fins configurations

2020 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Lei Luo ◽  
Yifeng Zhang ◽  
Chenglong Wang ◽  
Songtao Wang ◽  
Bengt Ake Sunden
Keyword(s):  
Heat Transfer ◽  
Leading Edge ◽  
Target Surface ◽  
Edge Region ◽  
Heat Transfer Characteristics ◽  
Content Type ◽  
Pin Fin ◽  
Transfer Characteristics ◽  
Turbine Vane ◽  
Pin Fins

Purpose The pin fin is applied into a Lamilloy cooling structure which is broadly used in the leading edge region of the modern gas turbine vane. The purpose of this paper is to investigate effects of the layout, diameter and shape of pin fins on the flow structure and heat transfer characteristics in a newly improved Lamilloy structure at the leading edge region of a turbine vane. Design/methodology/approach A numerical method is applied to investigate effects of the layout, diameter and shape of pin fins on the flow structure and heat transfer characteristics in a newly improved Lamilloy structure at the leading edge of a turbine vane. The diverse locations of pin fins are Lp = 0.35, 0.5, 0.65. The diameter of the pin fins varies from 8 mm to 32 mm. Three different ratios of root to roof diameter for pin fins are also investigated, i.e. k = 0.5, 1, 2. The Reynolds number ranges from 10,000 and 50,000. Results of the flow structures, heat transfer on the target surface and pin fin surfaces, and friction factor are studied. Findings The heat transfer on the pin fin surface gradually decreases and then increases as the location of the pin fins increases. Increasing the diameter of the pin fins causes the heat transfer on the pin fin surface to gradually increase, while a lower value of the friction factor occurs. Besides, the heat transfer on the pin fin surface at a small root diameter increases remarkably, but a slight heat transfer penalty is found at the target surface. It is also found that both the Reynolds analogy performance and the thermal performance are increased compared to the baseline whose diameter and normalized location of pin fins are set as 16 and 0.5 mm, respectively. Social implications The models provide a basic theoretical study to deal with nonuniformity of the temperature field for the turbine vane leading edge. The investigation also provides a better understanding of the heat transfer and flow characteristics in the leading edge region of a modern turbine vane. Originality/value This is a novel method to adopt pin fins into a Lamilloy cooling structure with curvature. It presents that the heat transfer of the pin fin surface in a pin-fin Lamilloy cooling structure with curvature can be significantly increased by changing the parameters of the pin fins which may lead to various flow behavior. In addition, the shape of the pin fin also shows great influence on the heat transfer and flow characteristics. However, the heat transfer of the target surface shows a small sensitivity to different layouts, diameter and shape of pin fin.

Heat Transfer and Flowfield Measurements in the Leading Edge Region of a Stator Vane Endwall

1999 ◽  
Vol 121 (3) ◽  
pp. 558-568 ◽  
Author(s):  
M. B. Kang ◽  
A. Kohli ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Secondary Flows ◽  
Edge Region ◽  
Leading Edge Vortex ◽  
Stator Vane ◽  
Edge Vortex ◽  
The Mean

The leading edge region of a first-stage stator vane experiences high heat transfer rates, especially near the endwall, making it very important to get a better understanding of the formation of the leading edge vortex. In order to improve numerical predictions of the complex endwall flow, benchmark quality experimental data are required. To this purpose, this study documents the endwall heat transfer and static pressure coefficient distribution of a modern stator vane for two different exit Reynolds numbers (Reex = 6 × 105 and 1.2 × 106). In addition, laser-Doppler velocimeter measurements of all three components of the mean and fluctuating velocities are presented for a plane in the leading edge region. Results indicate that the endwall heat transfer, pressure distribution, and flowfield characteristics change with Reynolds number. The endwall pressure distributions show that lower pressure coefficients occur at higher Reynolds numbers due to secondary flows. The stronger secondary flows cause enhanced heat transfer near the trailing edge of the vane at the higher Reynolds number. On the other hand, the mean velocity, turbulent kinetic energy, and vorticity results indicate that leading edge vortex is stronger and more turbulent at the lower Reynolds number. The Reynolds number also has an effect on the location of the separation point, which moves closer to the stator vane at lower Reynolds numbers.

scholarly journals Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

Endwall Heat Transfer Measurements in a Staggered Pin Fin Array With an Adiabatic Pin

2007 ◽  
Author(s):  
Forrest E. Ames ◽  
Chad A. Nordquist ◽  
Lindsay A. Klennert
Keyword(s):  
Heat Transfer ◽  
Turbulent Transport ◽  
Infrared Camera ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Skin Friction Coefficient ◽  
Vortex System ◽  
Flow Acceleration ◽  
Pin Fin ◽  
Pin Fin Array

Full surface endwall heat transfer distributions have been acquired in a staggered pin fin array with the use of an infrared camera. Values are presented at Reynolds numbers of 3000, 10,000 and 30,000 based on pin diameter and average velocity through adjacent pins. Average endwall Nusselt numbers agree closely with archival values at each Reynolds number. Locally averaged heat transfer levels show a substantial increase from the inlet through the first few rows and finally a nearly streamwise periodic condition in the second half of the eight row geometry. Increasing levels of heat transfer in the inlet region can be attributed to the leading edge vortex system, flow acceleration around pins, and the generation of turbulence. Distributions of turbulence intensity and turbulent scale are shown to help document the turbulent transport conditions through the array. Detailed endwall Nusselt number distributions are presented and compared at the three Reynolds numbers for the first four and last four rows. These detailed heat transfer distributions highlight the influence of the horseshoe vortex system in the entrance region and the wake generated turbulence throughout the pin fin array. Local velocity and turbulence distributions are presented together with local Stanton number and skin friction coefficient data to examine the aggressive nature of the turbulent mixing.

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