A Converging Slot-Hole Film-Cooling Geometry—Part 1: Low-Speed Flat-Plate Heat Transfer and Loss

2002 ◽  
Vol 124 (3) ◽  
pp. 453-460 ◽  
Author(s):  
J. E. Sargison ◽  
S. M. Guo ◽  
M. L. G. Oldfield ◽  
G. D. Lock ◽  
A. J. Rawlinson
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Flat Plate ◽  
Film Cooling ◽  
Large Scale ◽  
Guide Vane ◽  
Constant Heat Flux ◽  
Low Speed ◽  
Aerodynamic Loss ◽  
Cooling Hole

This paper presents experimental measurements of the performance of a new film-cooling hole geometry—the con¯vergings¯lot-hole¯ or console. This novel, patented geometry has been designed to improve the heat transfer and aerodynamic loss performance of turbine vane and rotor blade cooling systems. The physical principles embodied in the new hole design are described, and a typical example of the console geometry is presented. The cooling performance of a single row of consoles was compared experimentally with that of typical 35-deg cylindrical and fan-shaped holes and a slot, on a large-scale, flat-plate model at engine representative Reynolds numbers in a low-speed tunnel with ambient temperature main flow. The hole throat area per unit width is matched for all four hole geometries. By independently varying the temperature of the heated coolant and the heat flux from an electrically heated, thermally insulated, constant heat flux surface, both the heat transfer coefficient and the adiabatic cooling effectiveness were deduced from digital photographs of the color play of narrow-band thermochromic liquid crystals on the model surface. A comparative measurement of the aerodynamic losses associated with each of the four film-cooling geometries was made by traversing the boundary layer at the downstream end of the flat plate. The promising heat transfer and aerodynamic performance of the console geometry have justified further experiments on an engine representative nozzle guide vane in a transonic annular cascade presented in Part 2 of this paper.

A Converging Slot-Hole Film-Cooling Geometry: Part 1 — Low-Speed Flat-Plate Heat Transfer and Loss

2001 ◽  
Author(s):  
J. E. Sargison ◽  
S. M. Guo ◽  
M. L. G. Oldfield ◽  
G. D. Lock ◽  
A. J. Rawlinson
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Flat Plate ◽  
Film Cooling ◽  
Large Scale ◽  
Guide Vane ◽  
Constant Heat Flux ◽  
Low Speed ◽  
Aerodynamic Loss ◽  
Cooling Hole

This paper presents experimental measurements of the performance of a new film cooling hole geometry - the Converging Slot-Hole or Console. This novel, patented geometry has been designed to improve the heat transfer and aerodynamic loss performance of turbine vane and rotor blade cooling systems. The physical principles embodied in the new hole design are described, and a typical example of the console geometry is presented. The cooling performance of a single row of consoles was compared experimentally with that of typical 35° cylindrical and fan-shaped holes and a slot, on a large-scale, flat-plate model at engine representative Reynolds numbers in a low speed tunnel with ambient temperature main flow. The hole throat area per unit width is matched for all four hole geometries. By independently varying the temperature of the heated coolant and the heat flux from an electrically heated, thermally insulated, constant heat flux surface, both the heat transfer coefficient and the adiabatic cooling effectiveness were deduced from digital photographs of the colour play of narrow-band thermochromic liquid crystals on the model surface. A comparative measurement of the aerodynamic losses associated with each of the four film-cooling geometries was made by traversing the boundary layer at the downstream end of the flat plate. The promising heat transfer and aerodynamic performance of the console geometry have justified further experiments on an engine representative nozzle guide vane in a transonic annular cascade presented in Part 2 of this paper [1].

Heat Transfer Coefficient Augmentation for a Shaped Film Cooling Hole at a Range of Compound Angles

2020 ◽  
pp. 1-30
Author(s):  
Shane Haydt ◽  
Stephen Lynch
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Constant Heat Flux ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Compound Angle ◽  
The Heat Transfer Coefficient

Abstract Film cooling holes with shaped diffusers are used to efficiently deliver coolant to the surface of a gas turbine part to keep metal temperatures low. Reducing the heat flux into a component, relative to a case with no coolant injection, is the ultimate goal of film cooling. This reduction in heat flux is primarily achieved via a lower driving temperature at the wall for convection, represented by the adiabatic effectiveness. Another important consideration, however, is how the disturbance to the flowfield and thermal field caused by the injection of coolant augments the heat transfer coefficient. The present study examines the spatially-resolved heat transfer coefficient augmentation, measured using a constant heat flux foil and IR thermography, for a shaped film cooling hole at a range of compound angles. Results show that the heat transfer coefficient increases with compound angle and with blowing ratio. Due to the unique asymmetric flowfield of a compound angle hole, a significant amount of augmentation occurs to the side of the film cooling jet, where very little coolant is present. This causes local regions of increased heat flux, which is counter to the goal of film cooling. Heat transfer results are compared with adiabatic effectiveness and flowfield measurements from a previous study.

Heat Transfer Coefficient Augmentation for a Shaped Film Cooling Hole at a Range of Compound Angles

2019 ◽  
Author(s):  
Shane Haydt ◽  
Stephen Lynch
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Constant Heat Flux ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Compound Angle ◽  
The Heat Transfer Coefficient

Abstract Shaped film cooling holes are used to efficiently deliver coolant to the surface of a gas turbine part to keep metal temperatures low. The ultimate goal of film cooling is to reduce the heat flux into a component, relative to a case with no coolant injection. This reduction in heat flux is primarily achieved via a lower driving temperature at the wall for convection, represented by the adiabatic effectiveness. Another important consideration, however, is how the disturbance to the flowfield and thermal field caused by the injection of coolant augments the heat transfer coefficient. The present study examines the spatially-resolved heat transfer coefficient augmentation for a shaped film cooling hole at a range of compound angles, using a constant heat flux foil and IR thermography. Results show that the heat transfer coefficient increases with compound angle and with blowing ratio. Due to the unique asymmetric flowfield of a compound angle hole, a significant amount of augmentation occurs to the side of the film cooling jet, where very little coolant is present. This causes local regions of increased heat flux, which is counter to the goal of film cooling. Heat transfer results are compared with adiabatic effectiveness and flowfield measurements from a previous study.

A Novel Antivortex Turbine Film-Cooling Hole Concept

2008 ◽  
Vol 130 (3) ◽  
Author(s):  
James D. Heidmann ◽  
Srinath Ekkad
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Circular Cross Section ◽  
Round Hole ◽  
Net Heat Flux ◽  
Cooling Hole ◽  
Film Cooling Holes

A novel turbine film-cooling hole shape has been conceived and designed at NASA Glenn Research Center. This “antivortex” design is unique in that it requires only easily machinable round holes, unlike shaped film-cooling holes and other advanced concepts. The hole design is intended to counteract the detrimental vorticity associated with standard circular cross-section film-cooling holes. This vorticity typically entrains hot freestream gas and is associated with jet separation from the turbine blade surface. The antivortex film-cooling hole concept has been modeled computationally for a single row of 30 deg angled holes on a flat surface using the 3D Navier–Stokes solver GLENN-HT. A blowing ratio of 1.0 and density ratios of 1.05 and 2.0 are studied. Both film effectiveness and heat transfer coefficient values are computed and compared to standard round hole cases for the same blowing rates. A net heat flux reduction is also determined using both the film effectiveness and heat transfer coefficient values to ascertain the overall effectiveness of the concept. An improvement in film effectiveness of about 0.2 and in net heat flux reduction of about 0.2 is demonstrated for the antivortex concept compared to the standard round hole for both blowing ratios. Detailed flow visualization shows that as expected, the design counteracts the detrimental vorticity of the round hole flow, allowing it to remain attached to the surface.

Aerodynamics of a Letterbox Trailing Edge: Effects of Blowing Rate, Reynolds Number, and External Turbulence on Aerodynamic Losses and Pressure Distribution

2010 ◽  
Vol 132 (4) ◽  
Author(s):  
N. J. Fiala ◽  
J. D. Johnson ◽  
F. E. Ames
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Film Cooling ◽  
Large Scale ◽  
Reynolds Numbers ◽  
Design Flow ◽  
Trailing Edge ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
External Turbulence

A letterbox trailing edge configuration is formed by adding flow partitions to a gill slot or pressure side cutback. Letterbox partitions are a common trailing edge configuration for vanes and blades, and the aerodynamics of these configurations are consequently of interest. Exit surveys detailing total pressure loss, turning angle, and secondary velocities have been acquired for a vane with letterbox partitions in a large-scale low speed cascade facility. These measurements are compared with exit surveys of both the base (solid) and gill slot vane configurations. Exit surveys have been taken over a four to one range in chord Reynolds numbers (500,000, 1,000,000, and 2,000,000) based on exit conditions and for low (0.7%), grid (8.5%), and aerocombustor (13.5%) turbulence conditions with varying blowing rate (50%, 100%, 150%, and 200% design flow). Exit loss, angle, and secondary velocity measurements were acquired in the facility using a five-hole cone probe at a measuring station representing an axial chord spacing of 0.25 from the vane trailing edge plane. Differences between losses with the base vane, gill slot vane, and letterbox vane for a given turbulence condition and Reynolds number are compared providing evidence of coolant ejection losses, and losses due to the separation off the exit slot lip and partitions. Additionally, differences in the level of losses, distribution of losses, and secondary flow vectors are presented for the different turbulence conditions at the different Reynolds numbers. The letterbox configuration has been found to have slightly reduced losses at a given flow rate compared with the gill slot. However, the letterbox requires an increased pressure drop for the same ejection flow. The present paper together with a related paper (2008, “Letterbox Trailing Edge Heat Transfer—Effects of Blowing Rate, Reynolds Number, and External Turbulence on Heat Transfer and Film Cooling Effectiveness,” ASME, Paper No. GT2008-50474), which documents letterbox heat transfer, is intended to provide designers with aerodynamic loss and heat transfer information needed for design evaluation and comparison with competing trailing edge designs.

Influence of Film Cooling Hole Angles and Geometries on Aerodynamic Loss and Net Heat Flux Reduction

2011 ◽  
Author(s):  
Chia Hui Lim ◽  
Graham Pullan ◽  
Peter Ireland
Keyword(s):  
Heat Flux ◽  
Film Cooling ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Design Parameters ◽  
Aerodynamic Loss ◽  
Design Engineers ◽  
Net Heat Flux ◽  
Cooling Hole ◽  
Shaped Hole

Turbine design engineers have to ensure that film cooling can provide sufficient protection to turbine blades from the hot mainstream gas, while keeping the losses low. Film cooling hole design parameters include inclination angle (α), compound angle (β), hole inlet geometry and hole exit geometry. The influence of these parameters on aerodynamic loss and net heat flux reduction is investigated, with loss being the primary focus. Low-speed flat plate experiments have been conducted at momentum flux ratios of IR = 0.16, 0.64 and 1.44. The film cooling aerodynamic mixing loss, generated by the mixing of mainstream and coolant, can be quantified using a three-dimensional analytical model that has been previously reported by the authors. The model suggests that for the same flow conditions, the aerodynamic mixing loss is the same for holes with different α and β but with the same angle between the mainstream and coolant flow directions (angle κ). This relationship is assessed through experiments by testing two sets of cylindrical holes with different α and β: one set with κ = 35°, another set with κ = 60°. The data confirm the stated relationship between α, β, κ and the aerodynamic mixing loss. The results show that the designer should minimise κ to obtain the lowest loss, but maximise β to achieve the best heat transfer performance. A suggestion on improving the loss model is also given. Five different hole geometries (α = 35.0°, β = 0°) were also tested: cylindrical hole, trenched hole, fan-shaped hole, D-Fan and SD-Fan. The D-Fan and the SD-Fan have similar hole exits to the fan-shaped hole but their hole inlets are laterally expanded. The external mixing loss and the loss generated inside the hole are compared. It was found that the D-Fan and the SD-Fan have the lowest loss. This is attributed to their laterally expanded hole inlets, which lead to significant reduction in the loss generated inside the holes. As a result, the loss of these geometries is ≈ 50% of the loss of the fan-shaped hole at IR = 0.64 and 1.44.

A Novel Anti-Vortex Turbine Film Cooling Hole Concept

2007 ◽  
Author(s):  
James D. Heidmann ◽  
Srinath Ekkad
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Circular Cross Section ◽  
Round Hole ◽  
Net Heat Flux ◽  
Cooling Hole ◽  
Film Cooling Holes

A novel turbine film cooling hole shape has been conceived and designed at NASA Glenn Research Center. This “anti-vortex” design is unique in that it requires only easily machinable round holes, unlike shaped film cooling holes and other advanced concepts. The hole design is intended to counteract the detrimental vorticity associated with standard circular cross-section film cooling holes. This vorticity typically entrains hot freestream gas and is associated with jet separation from the turbine blade surface. The anti-vortex film cooling hole concept has been modeled computationally for a single row of 30 degree angled holes on a flat surface using the 3D Navier-Stokes solver Glenn-HT. A blowing ratio of 1.0 and density ratios of 1.05 and 2.0 are studied. Both film effectiveness and heat transfer coefficient values are computed and compared to standard round hole cases for the same blowing rates. A net heat flux reduction is also determined using both the film effectiveness and heat transfer coefficient values to ascertain the overall effectiveness of the concept. An improvement in film effectiveness of about 0.2 and in net heat flux reduction of about 0.2 is demonstrated for the anti-vortex concept compared to the standard round hole for both blowing ratios. Detailed flow visualization shows that as expected, the design counteracts the detrimental vorticity of the round hole flow, allowing it to remain attached to the surface.

A Converging Slot-Hole Film-Cooling Geometry—Part 2: Transonic Nozzle Guide Vane Heat Transfer and Loss

2002 ◽  
Vol 124 (3) ◽  
pp. 461-471 ◽  
Author(s):  
J. E. Sargison ◽  
S. M. Guo ◽  
M. L. G. Oldfield ◽  
G. D. Lock ◽  
A. J. Rawlinson
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Wide Band ◽  
Nozzle Guide Vane ◽  
Cooling Hole ◽  
Hole Geometry ◽  
Temperature Levels ◽  
Nozzle Guide

This paper presents the first experimental measurements on an engine representative nozzle guide vane, of a new film-cooling hole geometry, a con¯vergings¯lot-hole¯ or console. The patented console geometry is designed to improve the heat transfer and aerodynamic performance of turbine vane and rotor blade cooling systems. These experiments follow the successful validation of the console design in low-speed flat-plate tests described in Part 1 of this paper. Stereolithography was used to manufacture a resin model of a transonic, engine representative nozzle guide vane in which seven rows of previously tested fan-shaped film-cooling holes were replaced by four rows of consoles. This vane was mounted in the annular vane ring of the Oxford cold heat transfer tunnel for testing at engine Reynolds numbers, Mach numbers and coolant to mainstream momentum flux ratios using a heavy gas to simulate the correct coolant to mainstream density ratio. Heat transfer data were measured using wide-band thermochromic liquid crystals and a modified analysis technique. Both surface heat transfer coefficient and the adiabatic cooling effectiveness were derived from computer-video records of hue changes during the transient tunnel run. The cooling performance, quantified by the heat flux at engine temperature levels, of the console vane compares favourably with that of the previously tested vane with fan-shaped holes. The new console film-cooling hole geometry offers advantages to the engine designer due to a superior aerodynamic efficiency over the fan-shaped hole geometry. These efficiency measurements are demonstrated by results from midspan traverses of a four-hole pyramid probe downstream of the nozzle guide vane.

Detailed Measurements of Local Heat Transfer Coefficient in the Entrance to Normal and Inclined Film Cooling Holes

1994 ◽  
Author(s):  
David R. H. Gillespie ◽  
Aaron R. Byerley ◽  
Peter T. Ireland ◽  
Zuolan Wang ◽  
Terry V. Jones ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Film Cooling ◽  
Large Scale ◽  
Temperature Sensitive ◽  
Gas Temperature ◽  
Entry Length ◽  
Cooling Hole ◽  
Cooling Passage ◽  
Film Cooling Holes

The local heal transfer inside the entrance to large scale models of film cooling holes has been measured using the transient heat transfer technique. The method employs temperature sensitive liquid crystals to measure the surface temperature of large scale perspex models. Full distributions of local Nusselt number were calculated based on the cooling passage centreline gas temperature ahead of the cooling hole. The circumferentially averaged Nusselt number was also calculated based on the local mixed bulk driving gas temperature to aid interpretation of the results, and to broaden the potential application of the data. Data are presented for a single film cooling hole inclined at 90 and 150 degrees to the coolant duct wall. Both holes exhibited entry length heat transfer levels which were significantly lower than those predicted by entry length data in the presence of crossflow. The reasons for the comparative reduction are discussed in terms of the interpreted flow field.

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