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.

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

2013 ◽  
Vol 135 (5) ◽  
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 deg, and another set with κ = 60 deg. The data confirm the stated relationship between α, β, κ and the aerodynamic mixing loss. The results show that the designer should minimize κ to obtain the lowest loss, but maximize 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 deg, β = 0 deg) 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 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.

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 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.

Performance Evaluation of a Novel Film-Cooling Hole

2012 ◽  
Vol 134 (10) ◽  
Author(s):  
Ki-Don Lee ◽  
Kwang-Yong Kim
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Three Dimensional ◽  
Lateral Spreading ◽  
The Novel ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Shaped Hole

This paper presents a numerical investigation of the film-cooling performance of a novel film-cooling hole in comparison with a fan-shaped hole. The novel shaped hole is designed to increase the lateral spreading of coolant on the cooling surface. The film-cooling performance of the novel shaped hole is evaluated at a density ratio of 1.75 and the range of the blowing ratio of 0.5–2.5. The simulations were performed using three-dimensional Reynolds-averaged Navier–Stokes analysis with the SST k-ω model. The numerical results for the fan-shaped hole show very good agreement with the experimental data. For the blowing ratio of 0.5, the novel shaped film-cooling hole shows a similar cooling performance as the fan-shaped hole. However, as the blowing ratio increases, the novel shaped hole shows greatly improved lateral spreading of the coolant and the cooling performance in terms of the film-cooling effectiveness in comparison with the fan-shaped hole.

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].

Film Cooling Hole Shape Optimization Using Proper Orthogonal Decomposition

2014 ◽  
Author(s):  
Kozo Nita ◽  
Yoji Okita ◽  
Chiyuki Nakamata ◽  
Seiji Kubo ◽  
Kazuo Yonekura ◽  
...  
Keyword(s):  
Proper Orthogonal Decomposition ◽  
Film Cooling ◽  
Turbine Blades ◽  
Orthogonal Decomposition ◽  
Optimum Shape ◽  
Design Parameters ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Proper Orthogonal

Film cooling is a very effective cooling method for protecting the turbine blades exposed to hot gas from the heat. Since its cooling effectiveness is highly dependent on the shape of the hole, a wide variety of concepts and design parameters regarding hole shapes have been researched. However, there are no well-defined ways to determine the optimum shape of a film cooling hole. The CFD is a powerful tool for film cooling hole optimization. But with the number of parameters that define the film cooling hole shapes being so numerous, analytical optimization with CFD often requires computational resources that are unrealistic for the average design environment. Accordingly, for CFD to be effective in the optimization process, it is necessary to reduce the number of computations or shorten the calculation time per computation. In order to solve this problem, this paper presents a novel approach of applying 3D-POD (3D-Proper Orthogonal Decomposition) to the optimization of film cooling holes. POD is one of the most important component analysis methods and has the potential to reduce the number of parameters. From the computation results, a solution group was made by the RSM (Response Surface Method) and assessment functions, i.e., film cooling effectiveness, heat transfer coefficient, mixing loss, concentration of stress and robustness were considered first. In the end, however, considering the sensitivity of each objective function, the optimal hole shapes were obtained with only the film effectiveness being evaluated. In the following sections, this method and its results are described in detail.

scholarly journals Comparison of Two-Equation Turbulence Models for Prediction of Heat Transfer on Film-Cooled Turbine Blades

1997 ◽  
Author(s):  
Vijay K. Garg ◽  
Ali A. Ameri
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Turbulence Models ◽  
Computer Time ◽  
Suction Surface ◽  
The Heat Transfer Coefficient

A three-dimensional Navier-Stokes code has been used to compute the heat transfer coefficient on two film-cooled turbine blades, namely the VKI rotor with six rows of cooling holes including three rows on the shower head, and the C3X vane with nine rows of holes including five rows on the shower head. Predictions of heat transfer coefficient at the blade surface using three two-equation turbulence models, specifically, Coakley’s q-ω model, Chien’s k-ε model and Wilcox’s k-ω model with Menter’s modifications, have been compared with the experimental data of Camci and Arts (1990) for the VKI rotor, and of Hylton et al. (1988) for the C3X vane along with predictions using the Baldwin-Lomax (B-L) model taken from Garg and Gaugler (1995). It is found that for the cases considered here the two-equation models predict the blade heat transfer somewhat better than the B-L model except immediately downstream of the film-cooling holes on the suction surface of the VKI rotor, and over most of the suction surface of the C3X vane. However, all two-equation models require 40% more computer core than the B-L model for solution, and while the q-ω and k-ε models need 40% more computer time than the B-L model, the k-ω model requires at least 65% more time due to slower rate of convergence. It is found that the heat transfer coefficient exhibits a strong spanwise as well as streamwise variation for both blades and all turbulence models.

scholarly journals The Design of an Improved Endwall Film-Cooling Configuration

1998 ◽  
Author(s):  
S. Friedrichs ◽  
H. P. Hodson ◽  
W. N. Dawes
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Aerodynamic Loss ◽  
Cooling Flow ◽  
High Static Pressure ◽  
Endwall Film Cooling ◽  
Aerodynamic Losses

The endwall film-cooling cooling configuration investigated by Friedrichs et al. (1996, 1997) had in principle sufficient cooling flow for the endwall, but in practice, the redistribution of this coolant by secondary flows left large endwall areas uncooled. This paper describes the attempt to improve upon this datum cooling configuration by redistributing the available coolant to provide a better coolant coverage on the endwall surface, whilst keeping the associated aerodynamic losses small. The design of the new, improved cooling configuration was based on the understanding of endwall film-cooling described by Friedrichs et al. (1996, 1997). Computational fluid dynamics were used to predict the basic flow and pressure field without coolant ejection. Using this as a basis, the above described understanding was used to place cooling holes so that they would provide the necessary cooling coverage at minimal aerodynamic penalty. The simple analytical modelling developed in Friedrichs et al. (1997) was then used to check that the coolant consumption and the increase in aerodynamic loss lay within the limits of the design goal. The improved cooling configuration was tested experimentally in a large scale, low speed linear cascade. An analysis of the results shows that the redesign of the cooling configuration has been successful in achieving an improved coolant coverage with lower aerodynamic losses, whilst using the same amount of coolant as in the datum cooling configuration. The improved cooling configuration has reconfirmed conclusions from Friedrichs et al. (1996, 1997); firstly, coolant ejection downstream of the three-dimensional separation lines on the endwall does not change the secondary flow structures; secondly, placement of holes in regions of high static pressure helps reduce the aerodynamic penalties of platform coolant ejection; finally, taking account of secondary flow can improve the design of endwall film-cooling configurations.

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