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.

Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions: Part I—Impingement Cooling Without Film Coolant Extraction

1990 ◽  
Vol 112 (3) ◽  
pp. 451-458 ◽  
Author(s):  
R. S. Bunker ◽  
D. E. Metzger
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Nusselt Number ◽  
Internal Heat ◽  
Leading Edge ◽  
Diameter Ratio ◽  
Test Surface ◽  
Stagnation Line ◽  
Wide Range ◽  
Edge Sharpness

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. Cooling is provided by a single line of equally spaced multiple jets, aimed at the leading edge apex and exiting the leading edge region in the opposite or chordwise direction. 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 surface and are observed during the transient with automated computer vision and data acquisition systems. A wide range of parameter combinations of interest in cooled airfoil practice are 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 number and spanwise Nusselt number gradient as functions of chordwise position. Results indicate general increases in heat transfer with approximately the 0.6 power of jet Reynolds number, increases in heat transfer with decreasing leading edge sharpness as well as with decreasing nozzle-to-apex distance, and increases in spanwise average heat transfer with decreasing jet pitch-to-diameter ratio. The latter increases are accompanied by increases in the spanwise gradient of the heat transfer coefficient. Comparison with available prior results of much coarser spatial resolution shows good agreement and establishes confidence in the use of the results for design purposes and as baseline results for comparison with subsequent experiments involving film cooling bleed.

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.

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.

Wake-Induced Unsteady Stagnation-Region Heat Transfer Measurements

1994 ◽  
Vol 116 (1) ◽  
pp. 29-38 ◽  
Author(s):  
P. J. Magari ◽  
L. E. LaGraff
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Heat Transfer Rate ◽  
Transfer Rate ◽  
Edge Region ◽  
Isentropic Compression ◽  
Stagnation Region ◽  
Stagnation Line ◽  
Wide Range ◽  
Mach Numbers

An experimental investigation of wake-induced unsteady heat transfer in the stagnation region of a cylinder was conducted. The objective of the study was to create a quasi-steady representation of the stator/rotor interaction in a gas turbine using two stationary cylinders in crossflow. In this simulation, a larger cylinder, representing the leading-edge region of a rotor blade, was immersed in the wake of a smaller cylinder, representing the trailing-edge region of a stator vane. Time-averaged and time-resolved heat transfer results were obtained over a wide range of Reynolds number at two Mach numbers: one incompressible and one transonic. The tests were conducted at Reynolds numbers, Mach numbers, and gas-to-wall temperature ratios characteristic of turbine engine conditions in an isentropic compression-heated transient wind tunnel (LICH tube). The augmentation of the heat transfer in the stagnation region due to wake unsteadiness was documented by comparison with isolated cylinder tests. It was found that the time-averaged heat transfer rate at the stagnation line, expressed in terms of the Frossling number (Nu/Re), reached a maximum independent of the Reynolds number. The power spectra and cross-correlation of the heat transfer signals in the stagnation region revealed the importance of large vortical structures shed from the upstream wake generator. These structures caused large positive and negative excursions about the mean heat transfer rate in the stagnation region.

Internal and External Cooling of a Full Coverage Effusion Cooling Plate: Effects of Double Wall Cooling Configuration and Conditions

2017 ◽  
Author(s):  
Zhong Ren ◽  
Sneha Reddy Vanga ◽  
Nathan Rogers ◽  
Phil Ligrani ◽  
Keith Hollingsworth ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Film Cooling ◽  
Reynolds Numbers ◽  
Test Plate ◽  
Blowing Ratio ◽  
Spatially Resolved ◽  
Nusselt Numbers ◽  
Streamwise Location ◽  
Effusion Hole

The present study provides new heat transfer data for both the surfaces of the full coverage effusion cooling plate within a double wall cooling test facility. To produce the cooling stream, a cold-side cross-flow supply for the effusion hole array is employed. Also utilized is a unique mainstream mesh heater, which provides transient thermal boundary conditions, after mainstream flow conditions are established. For the effusion cooled surface, presented are spatially-resolved distributions of surface adiabatic film cooling effectiveness, and surface heat transfer coefficients (measured using infrared thermography). For the coolant side, presented are spatially-resolved distributions of surface Nusselt numbers (measured using liquid crystal thermography). Of interest are the effects of streamwise development, blowing ratio, and Reynolds number. Streamwise hole spacing and spanwise hole spacing (normalized by effusion hole diameter) on the effusion plate are 15 and 4, respectively. Effusion hole diameter is 6.35 mm, effusion hole angle is 25 degrees, and effusion plate thickness is 3 hole diameters. Considered are overall effusion blowing ratios from 2.9 to 7.5, with subsonic, incompressible flow, and constant freestream velocity with streamwise development, for two different mainstream Reynolds numbers. For the hot side (mainstream) of the effusion film cooling test plate, results for two mainflow Reynolds numbers of about 145000 and 96000 show that the adiabatic cooling effectiveness is generally higher for the lower Reynolds number for a particular streamwise location and blowing ratio. The heat transfer coefficient is generally higher for the low Reynolds number flow. This is due to altered supply passage flow behavior, which causes a reduction in coolant lift-off of the film from the surface as coolant momentum, relative to mainstream momentum, decreases. For the coolant side of the effusion test plate, Nusselt numbers generally increase with blowing ratio, when compared at a particular streamwise location and mainflow Reynolds number.

scholarly journals Highly Resolved Distribution of Heat Transfer for Turbine Leading Edge Film Cooling Including Reynolds Number and Blowing Rate Effects

1998 ◽  
Author(s):  
K. Jung ◽  
D. K. Hennecke
Keyword(s):  
Heat Transfer ◽  
Mass Transfer ◽  
Reynolds Number ◽  
Film Cooling ◽  
Heat Mass Transfer ◽  
Leading Edge ◽  
Transfer Coefficients ◽  
Complex Flow ◽  
Long Distance ◽  
Naphthalene Sublimation Technique

The effect of leading edge film cooling on heat transfer was experimentally investigated using the naphthalene sublimation technique. The experiments were performed on a symmetrical model of the leading edge suction side region of a high pressure turbine blade with one row of film cooling holes on each side. Two different lateral inclinations of the injection holes were studied: 0° and 45°. In order to build a data base for the validation and improvement of numerical computations, highly resolved distributions of the heat/mass transfer coefficients were measured. Reynolds numbers (based on hole diameter) were varied from 4000 to 8000 and blowing rate from 0.0 to 1.5. For better interpretation, the results were compared with injection-flow visualizations. Increasing the blowing rate causes more interaction between the jets and the mainstream, which creates higher jet turbulence at the exit of the holes resulting in a higher relative heat transfer. This increase remains constant over quite a long distance dependent on the Reynolds number. Increasing the Reynolds number keeps the jets closer to the wall resulting in higher relative heat transfer. The highly resolved heat/mass transfer distribution shows the influence of the complex flow field in the near hole region on the heat transfer values along the surface.

An Experimental Study of Turbine Vane Heat Transfer With Leading Edge and Downstream Film Cooling

1990 ◽  
Vol 112 (3) ◽  
pp. 477-487 ◽  
Author(s):  
N. V. Nirmalan ◽  
L. D. Hylton
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Leading Edge ◽  
External Heat ◽  
Gas Temperature ◽  
Temperature Ratio ◽  
External Heat Transfer ◽  
Turbine Vane

This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine vane external heat transfer. Steady-state experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters—Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio—were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The vane external heat transfer data obtained in this program indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil. The data presented in this paper illustrate the interaction of these variables and should provide the airfoil designer and computational analyst with the information required to improve heat transfer design capabilities for film-cooled turbine airfoils.

Spatially Resolved Heat Transfer and Friction Factors in a Rectangular Channel With 45-Deg Angled Crossed-Rib Turbulators

2003 ◽  
Vol 125 (3) ◽  
pp. 575-584 ◽  
Author(s):  
P. M. Ligrani ◽  
G. I. Mahmood
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Nusselt Number ◽  
Rectangular Channel ◽  
Test Surface ◽  
Spatially Resolved ◽  
Nusselt Numbers ◽  
Friction Factors ◽  
Smooth Channel ◽  
Rib Turbulators

Spatially resolved Nusselt numbers, spatially averaged Nusselt numbers, and friction factors are presented for a stationary channel with an aspect ratio of 4 and angled rib turbulators inclined at 45 deg with perpendicular orientations on two opposite surfaces. Results are given at different Reynolds numbers based on channel height from 10,000 to 83,700. The ratio of rib height to hydraulic diameter is .078, the rib pitch-to-height ratio is 10, and the blockage provided by the ribs is 25% of the channel cross-sectional area. Nusselt numbers are given both with and without three-dimensional conduction considered within the acrylic test surface. In both cases, spatially resolved local Nusselt numbers are highest on tops of the rib turbulators, with lower magnitudes on flat surfaces between the ribs, where regions of flow separation and shear layer reattachment have pronounced influences on local surface heat transfer behavior. The augmented local and spatially averaged Nusselt number ratios (rib turbulator Nusselt numbers normalized by values measured in a smooth channel) vary locally on the rib tops as Reynolds number increases. Nusselt number ratios decrease on the flat regions away from the ribs, especially at locations just downstream of the ribs, as Reynolds number increases. When adjusted to account for conduction along and within the test surface, Nusselt number ratios show different quantitative variations (with location along the test surface), compared to variations when no conduction is included. Changes include: (i) decreased local Nusselt number ratios along the central part of each rib top surface as heat transfer from the sides of each rib becomes larger, and (ii) Nusselt number ratio decreases near corners, where each rib joins the flat part of the test surface, especially on the downstream side of each rib. With no conduction along and within the test surface (and variable heat flux assumed into the air stream), globally-averaged Nusselt number ratios vary from 2.92 to 1.64 as Reynolds number increases from 10,000 to 83,700. Corresponding thermal performance parameters also decrease as Reynolds number increases over this range, with values in approximate agreement with data measured by other investigators in a square channel also with 45 deg oriented ribs.

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