Experimental Investigation on the Heat Transfer of a Leading Edge Cooling System: Effects of Jet-to-Jet Spacing and Showerhead Extraction

2013 ◽  
Author(s):  
Bruno Facchini ◽  
Francesco Maiuolo ◽  
Lorenzo Tarchi ◽  
Nils Ohlendorf
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Large Scale ◽  
Cooling System ◽  
Leading Edge ◽  
Test Facility ◽  
Cooling Flow ◽  
Nusselt Numbers ◽  
Circular Holes ◽  
Nusselt Number Distribution

An experimental survey of a leading edge cooling scheme was performed to measure the Nusselt number distribution on a large scale test facility simulating the leading edge cavity of a pressure turbine blade. Test section is composed by two adjacent cavities, a rectangular supply channel and the leading edge cavity. The cooling flow impinges on the concave leading edge internal walls, by means of an impingement array located between the two cavities, and it is extracted through shower-head and film cooling holes. The impingement geometry is composed of a double array of circular holes. The aim of the present study is to point out the effects on the heat transfer coefficient of the radial jet pitch (y/d = 3 to 5) and the tangential jet pitch (x/d = 3 to 5). Moreover the influence of the shower-head extraction on the heat transfer distribution is investigated. Measurements were performed by means of a transient technique using narrow band Thermo-chromic Liquid Crystals (TLC). Jet Reynolds number was varied in order to cover the typical engine conditions of these cooling systems (Rej = 15000–45000). Results are reported in terms of detailed 2D maps, radial and tangential averaged Nusselt numbers.

scholarly journals Experimental Investigation of a Leading Edge Cooling System With Optimized Inclined Racetrack Holes

2014 ◽  
Author(s):  
Carlo Carcasci ◽  
Bruno Facchini ◽  
Lorenzo Tarchi ◽  
Nils Ohlendorf
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Cooling System ◽  
Leading Edge ◽  
Test Facility ◽  
Cooling Flow ◽  
Nusselt Numbers ◽  
Nusselt Number Distribution ◽  
Scale Test ◽  
Film Cooling Holes

An experimental survey of a leading edge cooling scheme was performed to measure the Nusselt number distribution on a large scale test facility simulating the leading edge cavity of an high pressure turbine blade. Test section is composed by two adjacent cavities, a rectangular supply channel and the leading edge cavity. The cooling flow impinges on the concave leading edge internal walls, by means of an impingement array located between the two cavities, and it is extracted through showerhead and film cooling holes. The impingement geometry is composed by a double array of circular or shaped holes. The aim of the present study is to investigate the heat transfer performance of two optimized impingement schemes in comparison with a standard one with circular and orthogonal holes. Both the optimized arrays have inclined racetrack shaped holes and one of them has also a converging shape. Measurements were performed by means of a transient technique using narrow band Thermo-chromic Liquid Crystals (TLC). Jet Reynolds number was varied in order to cover the typical engine conditions of these cooling systems (Rej = 15000–45000). Results are reported in terms of detailed 2D maps, radial and tangential averaged Nusselt numbers.

Heat Transfer Measurements in a Leading Edge Geometry With Racetrack Holes and Film Cooling Extraction

2012 ◽  
Author(s):  
Luca Andrei ◽  
Carlo Carcasci ◽  
Riccardo Da Soghe ◽  
Bruno Facchini ◽  
Francesco Maiuolo ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Mass Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Internal Heat ◽  
Leading Edge ◽  
Jet Impingement ◽  
Internal Surface ◽  
Test Facility ◽  
Supply Channel

An experimental survey on a state of the art leading edge cooling scheme was performed to evaluate heat transfer coefficients (HTC) on a large scale test facility simulating an high pressure turbine airfoil leading edge cavity. Test section includes a trapezoidal supply channel with three large racetrack impingement holes. On the internal surface of the leading edge, four big fins are placed in order to confine impingement jets. The coolant flow impacts the leading edge internal surface and it is extracted from the leading edge cavity through 24 showerhead holes and 24 film cooling holes. The aim of the present study is to investigate the combined effects of jet impingement and mass flow extraction on the internal heat transfer of the leading edge. A non uniform mass flow extraction was also imposed to reproduce the effects of pressure side and suction side external pressure. Measurements were performed by means of a transient technique using narrow band Thermo-chromic Liquid Crystals (TLC). Jet Reynolds number and crossflow conditions into the supply channel were varied in order to cover the typical engine conditions of these cooling systems (Rej = 10000–40000). Experiments were compared with a numerical analysis on the same test case in order to better understand flow interaction inside the cavity. Results are reported in terms of detailed 2D maps, radial-wise and span-wise averaged values of Nusselt number.

Heat Transfer Measurements in a Leading Edge Geometry With Racetrack Holes and Film Cooling Extraction

2013 ◽  
Vol 135 (3) ◽  
Author(s):  
Luca Andrei ◽  
Carlo Carcasci ◽  
Riccardo Da Soghe ◽  
Bruno Facchini ◽  
Francesco Maiuolo ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Mass Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Internal Heat ◽  
Leading Edge ◽  
Jet Impingement ◽  
Internal Surface ◽  
Test Facility ◽  
Supply Channel

An experimental survey on a state of the art leading edge cooling scheme was performed to evaluate heat transfer coefficients (HTC) on a large scale test facility simulating a high pressure turbine airfoil leading edge cavity. The test section includes a trapezoidal supply channel with three large racetrack impingement holes. On the internal surface of the leading edge, four big fins are placed in order to confine impingement jets. The coolant flow impacts the leading edge internal surface, and it is extracted from the leading edge cavity through 24 showerhead holes and 24 film cooling holes. The aim of the present study is to investigate the combined effects of jet impingement and mass flow extraction on the internal heat transfer of the leading edge. A nonuniform mass flow extraction was also imposed to reproduce the effects of the pressure side and suction side external pressure. Measurements were performed by means of a transient technique using narrow band thermochromic liquid crystals (TLCs). Jet Reynolds number and crossflow conditions into the supply channel were varied in order to cover the typical engine conditions of these cooling systems (Rej=10,000-40,000). Experiments were compared with a numerical analysis on the same test case in order to better understand flow interaction inside the cavity. Results are reported in terms of detailed 2D maps, radial-wise, and span-wise averaged values of Nusselt number.

Influence of Combustor Swirl on Endwall Heat Transfer and Film Cooling Effectiveness at the Large Scale Turbine Rig

2017 ◽  
Vol 139 (8) ◽  
Author(s):  
Holger Werschnik ◽  
Jonathan Hilgert ◽  
Manuel Wilhelm ◽  
Martin Bruschewski ◽  
Heinz-Peter Schiffer
Keyword(s):  
Heat Transfer ◽  
Mass Flux ◽  
Film Cooling ◽  
Large Scale ◽  
Guide Vane ◽  
Leading Edge ◽  
Double Row ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nusselt Numbers

At the large scale turbine rig (LSTR) at Technische Universität Darmstadt, Darmstadt, Germany, the aerothermal interaction of combustor exit flow conditions on the subsequent turbine stage is examined. The rig resembles a high pressure turbine and is scaled to low Mach numbers. A baseline configuration with an axial inflow and a swirling inflow representative for a lean combustor is modeled by swirl generators, whose clocking position toward the nozzle guide vane (NGV) leading edge can be varied. A staggered double-row of cylindrical film cooling holes on the endwall is examined. The effect of swirling inflow on heat transfer and film cooling effectiveness is studied, while the coolant mass flux rate is varied. Nusselt numbers are calculated using infrared thermography and the auxiliary wall method. Boundary layer, turbulence, and five-hole probe measurements as well as numerical simulations complement the examination. The results for swirling inflow show a decrease of film cooling effectiveness of up to 35% and an increase of Nusselt numbers of 10–20% in comparison to the baseline case for low coolant mass flux rates. For higher coolant injection, the heat transfer is on a similar level as the baseline. The differences vary depending on the clocking position. The turbulence intensity is increased to 30% for swirling inflow.

Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions: Part II—Impingement Cooling With Film Coolant Extraction

1990 ◽  
Vol 112 (3) ◽  
pp. 459-466 ◽  
Author(s):  
D. E. Metzger ◽  
R. S. Bunker
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Nusselt Number ◽  
Film Cooling ◽  
Large Scale ◽  
Internal Heat ◽  
Leading Edge ◽  
Stagnation Line ◽  
Nusselt Numbers ◽  
Wide Range

An experimental study has been designed and performed to measure very localized internal heat transfer characteristics in large-scale models of turbine blade impingement-cooled leading edge regions that allow extraction, or bleed-off, of a portion of the internal cooling flow to provide leading edge film cooling along the blade external surface. The internal impingement air is provided by a single line of equally spaced multiple jets, aimed at the leading edge apex and generally exiting, minus the bleed-off flow, in the opposite or chordwise direction. The film coolant flow extraction takes place through two lines of holes, one each on the blade suction side and the blade pressure side, both fairly close to the airfoil leading edge. Detailed two-dimensional local surface Nusselt number distributions have been obtained through the use of aerodynamically steady but thermally transient tests employing temperature-indicating coatings. The thin coatings are sprayed directly on the test surfaces, and are observed during a test transient with automated computer vision and data acquisition systems. A wide range of parameter combinations of interest in cooled airfoil practice is covered in the test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge sharpness, jet pitch-to-diameter ratio, and jet nozzle-to-apex travel distance. Measured local Nusselt numbers at each chordwise location back from the stagnation line have been used to calculate both the spanwise-average Nusselt numbers and spanwise Nusselt number gradients as functions of chordwise position. The results without film coolant extraction, presented in the Part I companion paper, are used as a basis of comparison to determine the additional effects of the film cooling bleed. Results indicate that heat transfer is primarily dependent on jet Reynolds number with smaller influences from the flow extraction rate. The results also suggest that changes in the spanwise alignment of the impingement nozzles relative to the position of the film cooling holes can cause significant variations in leading edge metal temperatures.

Clocking Effects of Inlet Non-Uniformities in a Fully Cooled High-Pressure Vane: A Conjugate Heat Transfer Analysis

2015 ◽  
Author(s):  
Duccio Griffini ◽  
Massimiliano Insinna ◽  
Simone Salvadori ◽  
Francesco Martelli
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Cooling System ◽  
Hot Spot ◽  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Heat Transfer Analysis

A high-pressure vane equipped with a realistic film-cooling configuration has been studied. The vane is characterized by the presence of multiple rows of fan-shaped holes along pressure and suction side while the leading edge is protected by a showerhead system of cylindrical holes. Steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) simulations have been performed. A preliminary grid sensitivity analysis with uniform inlet flow has been used to quantify the effect of spatial discretization. Turbulence model has been assessed in comparison with available experimental data. The effects of the relative alignment between combustion chamber and high-pressure vanes are then investigated considering realistic inflow conditions in terms of hot spot and swirl. The inlet profiles used are derived from the EU-funded project TATEF2. Two different clocking positions are considered: the first one where hot spot and swirl core are aligned with passage and the second one where they are aligned with the leading edge. Comparisons between metal temperature distributions obtained from conjugate heat transfer simulations are performed evidencing the role of swirl in determining both the hot streak trajectory within the passage and the coolant redistribution. The leading edge aligned configuration is resulted to be the most problematic in terms of thermal load, leading to increased average and local vane temperature peaks on both suction side and pressure side with respect to the passage aligned case. A strong sensitivity of both injected coolant mass flow and heat removed by heat sink effect has also been highlighted for the showerhead cooling system.

Numerical Investigation on the Effect of Rotation and Holes Arrangement in Cold Bridge-Type Impingement Cooling Systems

2020 ◽  
Vol 142 (4) ◽  
Author(s):  
Simone Paccati ◽  
Lorenzo Cocchi ◽  
Lorenzo Mazzei ◽  
Antonio Andreini
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Mutual Influence ◽  
Cooling System ◽  
Leading Edge ◽  
Lateral Spreading ◽  
Test Point ◽  
Navier Stokes ◽  
Specific Flow ◽  
Bridge Type

Abstract This work presents the results of a numerical analysis performed on a gas turbine leading edge cooling system. The investigation was carried out in order to provide a detailed interpretation of the outcomes of a parallel experimental campaign. The cooling geometry consists of a cold bridge-type impingement system: a radial channel feeds an array of holes, which in turn generate impingement jets cooling down the inner side of the leading edge surface. Coolant is extracted by five rows of holes, replicating film cooling and showerhead systems. Two impingement geometries were considered, presenting different holes arrangements and diameters but sharing the same overall passage area, in order to highlight the effect of different coolant distributions inside the leading edge cavity. For both geometries, a single test point was investigated in static and rotating conditions, with an equivalent slot Reynolds number of around 8200 and feeding conditions corresponding to the midspan radial section of the blade. Both steady Reynolds averaged Navier Stokes (RANS) approach and scale adaptive simulation (SAS) were tested. Due to the strong unsteadiness of the flow field, the latter proved to be superior: as a consequence, the SAS approach was adopted to study every case. A fairly good agreement was observed between the measured and computed heat transfer distributions, which allowed to exploit the numerical results to get a detailed description of the phenomena associated with the different cases. Results reveal that the two holes arrangements lead to strongly different heat transfer patterns, related to the specific flow phenomena occurring inside the leading edge cavity and to the mutual influence of the various system features. Rotational effects also appear to interact with the supply condition, altering the jet lateral spreading and the overall heat transfer performance.

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.

Aerodynamic and Heat Transfer Characterization of a Nozzle Vane Cascade With and Without Platform Cooling

2015 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Surface Flow ◽  
Secondary Flows ◽  
Intensity Level ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane

In the present paper, aerodynamic and thermal performance of a linear nozzle vane cascade is fully assessed. Tests have been carried out with and without platform cooling, with coolant ejected through a slot located upstream of the leading edge. Cooling air is also ejected through a row of cylindrical holes located upstream of the slot, simulating a combustor cooling system. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at variable cooling injection conditions. Aero-thermal characterization of vane platform was obtained through 5-hole probe measurements, oil & dye surface flow visualizations, measurements of end wall adiabatic film cooling effectiveness and heat transfer coefficient. The platform cooling scheme operated at nominal injection rate was shown to effectively reduce the heat load over most of the platform surface, with only a small increase in secondary flows loss. Combustor holes injection resulted beneficial in controlling momentum of coolant approaching the cascade, thus limiting the secondary flows growth and resulting in an increase of the coolant film length inside of the passage.

Endwall Effusion Cooling System Behaviour Within a High-Pressure Turbine Cascade: Part 2—Heat Transfer and Effectiveness Measurements

2010 ◽  
Author(s):  
B. Facchini ◽  
L. Tarchi ◽  
L. Toni ◽  
S. Zecchi
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Large Scale ◽  
Cooling System ◽  
Main Flow ◽  
Turbine Cascade ◽  
High Pressure Turbine ◽  
Cooling Flow ◽  
Micro Holes ◽  
Element Approach

The cooling performance of a micro-holed endwall of a large-scale high pressure turbine cascade has been investigated within the European Project AITEB-2. The experimental investigation has been performed for a baseline configuration, with a smooth solid endwall and with a micro-holed endwall providing micro-jets ejection from the wall. A micro-holed endwall made of two modules was adopted in order to reduce the compound angle between the main flow and the micro jets axes. The micro-holed endwall is provided with a total amount of 3294 micro-holes with a diameter of 0.1 per cent of the blade chord. Four different cooling flow rates, from 1.2% to 2.6% of the main flow mass flow rate respectively, were investigated and the experimental results are reported in the paper. Both adiabatic effectiveness and heat transfer coefficient have been measured employing a steady state technique with Thermo-chromic Liquid Crystals (TLC). A thin stainless steel heating foil was used to generate the surface heat flux for the HTC measurements and a data reduction procedure based on a Finite Element approach has been developed to take into account the non uniform heat generation along the endwall.

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