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.

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.

Numerical Analysis of Heat Transfer in a Leading Edge Geometry With Racetrack Holes and Film Cooling Extraction

2013 ◽  
Author(s):  
Luca Andrei ◽  
Antonio Andreini ◽  
Riccardo Da Soghe ◽  
Bruno Facchini ◽  
Stefano Zecchi
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Numerical Study ◽  
Internal Heat ◽  
Leading Edge ◽  
Jet Impingement ◽  
Internal Surface ◽  
Heat Transfer Process ◽  
University Of Florence ◽  
The Impact

A numerical study of a state of the art leading edge cooling scheme was performed to analyze the heat transfer process within the leading edge cavity of a high pressure turbine airfoil. The investigated geometries account a trapezoidal supply channel with a large racetrack impingement holes. The coolant jets, confined among two consequent large fins, impact the leading edge internal surface and it is extracted from the leading edge cavity through both showerhead holes and film cooling holes. The CFD setup has been validated by means of the experimental measurements performed on a dedicated test rig developed and operated at University of Florence. The aim of this study is to investigate the combined effects of jet impingement, mass flow extraction and fins presence on the internal heat transfer of the leading edge cavity. More in details, the paper analyses the impact, in terms of blade metal temperature, of large fins presence and positioning. Jet’s Reynolds number is varied in order to cover the typical engine conditions of these cooling systems (Rej = 20000 – 40000).

scholarly journals State-of-the-Art Cooling Technology for a Turbine Rotor Blade

2017 ◽  
Author(s):  
Jason Town ◽  
Doug Straub ◽  
James Black ◽  
Karen Thole ◽  
Tom Shih
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
State Of The Art ◽  
Internal Heat ◽  
Leading Edge ◽  
Jet Impingement ◽  
Test Facility ◽  
Turbine Wheel ◽  
Baseline Design ◽  
External Cooling

Effective internal and external cooling of airfoils is key to maintaining component life for efficient gas turbines. Cooling designs have spanned the range from simple internal convective channels to more advanced double-walls with shaped film-cooling holes. This paper describes the development of an internal and external cooling concept for a state-of-the-art cooled turbine blade. These cooling concepts are based on a review of literature and patents, as well as, interactions with academic and industry turbine cooling experts. The cooling configuration selected and described in this paper is referred to as the “baseline” design, since this design will simultaneously be tested with other more advanced blade cooling designs in a rotating turbine test facility using a “rainbow turbine wheel” configuration. For the baseline design, the leading edge is cooled by internal jet impingement and showerhead film cooling. The mid-chord region of the blade contains a three-pass serpentine passage with internal discrete V-shaped trip strips to enhance the internal heat transfer coefficient. The film cooling along the mid-chord of the blade uses multiple rows of shaped diffusion holes. The trailing edge is internally cooled using jet impingement and externally film cooled through partitioned cuts on the pressure side of the blade.

scholarly journals State-of-the-Art Cooling Technology for a Turbine Rotor Blade

2018 ◽  
Vol 140 (7) ◽  
Author(s):  
Jason Town ◽  
Douglas Straub ◽  
James Black ◽  
Karen A. Thole ◽  
Tom I-P. Shih
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
State Of The Art ◽  
Internal Heat ◽  
Leading Edge ◽  
Jet Impingement ◽  
Test Facility ◽  
Turbine Wheel ◽  
Baseline Design ◽  
External Cooling

Effective internal and external cooling of airfoils is key to maintaining component life for efficient gas turbines. Cooling designs have spanned the range from simple internal convective channels to more advanced double-walls with shaped film-cooling holes. This paper describes the development of an internal and external cooling concept for a state-of-the-art cooled turbine blade. These cooling concepts are based on a review of literature and patents, as well as, interactions with academic and industry turbine cooling experts. The cooling configuration selected and described in this paper is referred to as the “baseline” design, since this design will simultaneously be tested with other more advanced blade cooling designs in a rotating turbine test facility using a “rainbow turbine wheel” configuration. For the baseline design, the leading edge is cooled by internal jet impingement and showerhead film cooling. The midchord region of the blade contains a three-pass serpentine passage with internal discrete V-shaped trip strips to enhance the internal heat transfer coefficient (HTC). The film cooling along the midchord of the blade uses multiple rows of shaped diffusion holes. The trailing edge is internally cooled using jet impingement and externally film cooled through partitioned cuts on the pressure side of the blade.

An Internal Heat Transfer Study in a Cooled Nozzle Guide Vane of a Linear Cascade

2014 ◽  
Author(s):  
Arun Kumar Pujari ◽  
Prasad B. V. S. S. Subrahmanyaa ◽  
Sitaram Nekkanti
Keyword(s):  
Heat Transfer ◽  
Surface Temperature ◽  
Film Cooling ◽  
Guide Vane ◽  
Internal Heat ◽  
Leading Edge ◽  
Internal Surface ◽  
Structure Variation ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Experimental and computational heat transfer investigations are reported in the interior mid span of the pressure surface of a Nozzle Guide Vane (NGV) subjected to combined impingement and film cooling. The study is carried out by considering a two dimensional cascade domain having four passages formed between the five vane each has a chord length of 228 mm and spacing (between the blades) of 200 mm. The vane internal surface is cooled by two impingement inserts namely front and aft impingement tubes. The front impingement tube is used to cool the internal side of the leading edge of the NGV whereas the aft impingement tube is used to cool mainly the mid span of the internal surface. The mass flow through the impingement chamber is varied for a fixed target plate distance to jet diameter ratio of 1.12. The surface temperature at the mid chord region was measured by liquid crystal technique. The surface temperature obtained from both experiments and computations are compared and the computationally obtained average heat transfer coefficient distribution along chord reported. The flow structure variation along the chord and its effect on Nusselt number distribution is presented. The computation is carried out by using Shear stress transport (SST) k-ω turbulence model in the ANSY FLUENT-14 flow solver.

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.

Turbine Vane Leading Edge Impingement Cooling With a Sweeping Jet

2018 ◽  
Author(s):  
Lucas Agricola ◽  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Pressure Drop ◽  
Mass Flow ◽  
Guide Vane ◽  
Internal Heat ◽  
Leading Edge ◽  
Jet Impingement ◽  
Freestream Turbulence ◽  
Impingement Cooling ◽  
Jet Array

A low speed linear cascade was used to investigate sweeping jet impingement cooling in a nozzle guide vane leading edge at an engine-relevant Biot number. Sweeping and steady jets were studied at varying mass flow rates and freestream turbulence intensities. Infrared thermography and a thermal inertia technique were used to determine the overall cooling effectiveness and internal heat transfer coefficients of the impingement cooling configurations. The circular jet array provided higher overall effectiveness values at both freestream turbulence intensities. The sweeping jet array provided a broader heat transfer profile due to the spreading of the jet. Pressure drop was measured for each jet geometry, and the circular jet was found to have less pressure drop than the sweeping jet at a given mass flow rate.

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.

Experimental Evaluations of the Relative Contributions to Overall Effectiveness in Turbine Blade Leading Edge Cooling

2018 ◽  
Author(s):  
Carol E. Bryant ◽  
Connor J. Wiese ◽  
James L. Rutledge ◽  
Marc D. Polanka
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Leading Edge ◽  
Internal Surface ◽  
Internal Cooling ◽  
Transfer Coefficients ◽  
Impingement Cooling ◽  
External Cooling ◽  
Turbine Component ◽  
Overall Effectiveness

Gas turbine hot gas path components are protected through a combination of internal cooling and external film cooling. The coolant typically travels through internal passageways, which may involve impingement on the internal surface of a turbine component, before being ejected as film cooling. Internal cooling effects have been studied in facilities that allow measurement of heat transfer coefficients within models of the internal cooling paths, with large heat transfer coefficients generally desirable. External film cooling is typically evaluated through measurements of the adiabatic effectiveness and its effect on the external heat transfer coefficient. Efforts aimed at improving cooling are often focused on either only the internal cooling or the film cooling; however, the common coolant flow means the internal and external cooling schemes are linked and the coolant holes themselves provide another convective path for heat transfer to the coolant. Recently, measurements of overall cooling effectiveness using matched Biot number turbine component models allow evaluation of the nondimensional wall temperature achieved for the fully cooled component. However, the relative contributions of internal cooling, external cooling, and convection within the film cooling holes is not well understood. Large scale, matched Biot number experiments, complemented by CFD simulations, were performed on a fully film cooled cylindrical leading edge model to evaluate the effects of various alterations in the cooling design on the overall effectiveness. The relative influence of film cooling and cooling within the holes was evaluated by selectively disabling individual holes and quantifying how the overall effectiveness changed. Several internal impingement cooling schemes in addition to a baseline case without impingement cooling were also tested. In general, impingement cooling is shown to have a negligible influence on the overall effectiveness in the showerhead region. This indicates that the cost and pressure drop penalties for implementing impingement cooling may not be compensated by an increase in thermal performance. Instead, the internal cooling provided by convection within the holes themselves was shown, along with external film cooling, to be a dominant contribution to the overall cooling effectiveness. Indeed, the numerous holes within the showerhead region impede the ability of internal surface cooling schemes to influence the outside surface temperature. The results of this research may allow improved focus of future efforts on the forms of cooling with the greatest potential to improve cooling performance.

A Transient Calorimeter Technique for Determining Regional Heat Transfer Coefficients in the Three-Temperature Flowfield at a Turbine Airfoil Leading Edge

2005 ◽  
Author(s):  
Scott R. Nowlin ◽  
David R. H. Gillespie ◽  
Peter T. Ireland ◽  
Ralf Knoche ◽  
T. Robert Kingston
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Internal Flow ◽  
Leading Edge ◽  
Internal Surface ◽  
Reynolds Numbers ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Novel Method ◽  
Exposed Surface Area

In this paper, the authors develop a novel method of obtaining regionally-averaged heat transfer coefficients in flowfields characterized by three temperatures using the well-known transient calorimeter technique. The technique is used to determine heat transfer in aluminum models of idealized turbine blade leading edges cooled through internal surface impingement, film cooling feed passages, and external convective film cooling. The external surface is subject to a stagnating mainstream crossflow. Importantly, the contributions to heating from the external flow and cooling from the internal flow can be separately resolved solely by heating the internal flow. Results for a basic showerhead geometry and an advanced intersecting-passage cooling configuration are presented for a range of internal and external Reynolds numbers. The intersecting-passage model shows little improvement in heat transfer coefficient over the showerhead for the flow conditions tested; however, the total cooling carried out is improved by the increase in exposed surface area. The technique’s uncertainties are fully assessed.

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