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Relevant bibliographies by topics / Film cooling

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Academic literature on the topic 'Film cooling'

Author: Grafiati

Published: 4 June 2021

Last updated: 6 September 2023

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Journal articles on the topic "Film cooling"

1

Shi, Li, Zhiying Sun, and Yuanfeng Lu. "The Combined Influences of Film Cooling and Thermal Barrier Coatings on the Cooling Performances of a Film and Internal Cooled Vane." Coatings 10, no. 9 (September 5, 2020): 861. http://dx.doi.org/10.3390/coatings10090861.

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This paper presents a numerical investigation on the combined influences of film cooling and thermal barrier coatings (TBCs) on the cooling performances of a NASA C3X guide vane. The results show that: (1) film cooling on the pressure side is more effective than suction side, especially on the trailing edge where multiple cooling and thermal protection techniques include internal cooling and TBCs are necessary. (2) TBCs show positive and negative roles in improving cooling performance at the same time for the coated vane with or without film cooling. Without film cooling, TBCs show negative roles on the regions with lower temperature external hot gas, which is caused by flow acceleration from the stagnation line of the suction side. (3) Internal cooling improvement caused by coolant introduction leads to a larger cooling effectiveness inclement due to TBCs near coolant plenums and film cooling holes. However, the influence of TBCs on cooling effectiveness increment goes down and even shows negative roles on the regions away from coolant plenums and under the effective coverage of the film cooling. (4) Improving the convective heat transfer of coolant with the wall of coolant plenums and film cooling holes is the guarantee of improving the cooling performance of a coated vane.

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Harrington, Mark K., Marcus A. McWaters, David G. Bogard, Christopher A. Lemmon, and Karen A. Thole. "Full-Coverage Film Cooling With Short Normal Injection Holes." Journal of Turbomachinery 123, no. 4 (February 1, 2001): 798–805. http://dx.doi.org/10.1115/1.1400111.

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An experimental and computational investigation was conducted on the film cooling adiabatic effectiveness of a flat plate with full coverage film cooling. The full coverage film cooling array was comprised of ten rows of coolant holes, arranged in a staggered pattern, with short L/D=1, normal coolant holes. A single row of cooland holes was also examined to determine the accuracy of a superposition prediction of the full coverage adiabatic effectiveness performance. Large density coolant jets and high mainstream turbulence conditions were utilized to simulate realistic engine conditions. High-resolution adiabatic effectiveness measurements were obtained using infrared imaging techniques and a large-scale flat plate model. Optimum adiabatic effectiveness was found to occur for a blowing ratio of M=0.65. At this blowing ratio separation of the coolant jet immediately downstream of the hole was observed. For M=0.65, the high mainstream turbulence decreased the spatially averaged effectiveness level by 12 percent. The high mainstream turbulence produced a larger effect for lower blowing ratios. The superposition model based on single row effectiveness results over-predicted the full coverage effectiveness levels.

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Wilfert, Gu¨nter, and Stefan Wolff. "Influence of Internal Flow on Film Cooling Effectiveness." Journal of Turbomachinery 122, no. 2 (February 1, 1999): 327–33. http://dx.doi.org/10.1115/1.555449.

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Film cooling experiments were conducted to investigate the effects of internal flow conditions and plenum geometry on the film cooling effectiveness. The film cooling measurements show a strong influence of the coolant inlet conditions on film cooling performance. The present experiments were carried out on a flat plate with a row of cylindrical holes oriented at 30 deg with respect to a constant-velocity external flow, systematically varying the plenum geometry and blowing rates 0.5⩽M⩽1.25. Adiabatic film cooling measurements using the multiple narrow-banded thermochromic liquid crystal technique (TLC) were carried out, simulating a flow parallel to the mainstream flow with and without crossflow at the coolant hole entry compared with a standard plenum configuration. An impingement in front of the cooling hole entry with and without crossflow was also investigated. For all parallel flow configurations, ribs were installed at the top and bottom coolant channel wall. As the hole length-to-diameter ratio has an influence on the film cooling effectiveness, the wall thickness has also been varied. In order to optimize the benefit of the geometry effects with ribs, a vortex generator was designed and tested. Results from these experiments show in a region 5⩽X/D⩽80 downstream of the coolant injection location differences in adiabatic film cooling effectiveness between +5 percent and +65 percent compared with a standard plenum configuration. [S0889-504X(00)01102-8]

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4

Vasu Devan Nair Girija Kumari, Krishna Anand, and Parammasivam Kanjikoil Mahali. "Investigations of improved cooling effectiveness for ramp film cooling with compound angle film cooling jets." Aircraft Engineering and Aerospace Technology 93, no. 6 (July 28, 2021): 971–84. http://dx.doi.org/10.1108/aeat-05-2020-0082.

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Purpose This paper aims to investigate the film cooling effectiveness (FCE) and mixing flow characteristics of the flat surface ramp model integrated with a compound angled film cooling jet. Design/methodology/approach Three-dimensional numerical simulation is performed on a flat surface ramp model with Reynolds Averaged Navier-Stokes approach using a finite volume solver. The tested model has a fixed ramp angle of 24° and a ramp width of two times the diameter of the film cooling hole. The coolant air is injected at 30° along the freestream direction. Three different film hole compound angles oriented to freestream direction at 0°, 90° and 180° were investigated for their performance on-ramp film cooling. The tested blowing ratios (BRs) are in the range of 0.9–2.0. Findings The film hole oriented at a compound angle of 180° has improved the area-averaged FCE on the ramp test surface by 86.74% at a mid-BR of 1.4% and 318.75% at higher BRs of 2.0. The 180° film hole compound angle has also produced higher local and spanwise averaged FCE on the ramp test surface. Originality/value According to the authors’ knowledge, this study is the first of its kind to investigate the ramp film cooling with a compound angle film cooling hole. The improved ramp model with a 180° film hole compound angle can be effectively applied for the end-wall surfaces of gas turbine film cooling.

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Lutum, E., and B. V. Johnson. "Influence of the Hole Length-to-Diameter Ratio on Film Cooling With Cylindrical Holes." Journal of Turbomachinery 121, no. 2 (April 1, 1999): 209–16. http://dx.doi.org/10.1115/1.2841303.

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Film cooling experiments were conducted to investigate the effects of coolant hole length-to-diameter ratio on the film cooling effectiveness. The results from these experiments offer an explanation for the differences between the film cooling results for cylindrical hole injection configurations previously reported by Goldstein et al. (1974), Pedersen et al. (1977), and Sinha et al. (1991). The previously reported injection configurations differed primarily in coolant hole length-to-diameter ratio. The present experiments were conducted with a row of cylindrical holes oriented at 35 deg to a constant-velocity external flow, systematically varying the hole length-to-diameter ratios (L/D = 1.75, 3.5, 5, 7, and 18), and blowing rates (0.52 ≤ M ≤ 1.56). Results from these experiments show in a region 5 ≤ X/D ≤ 50 downstream of coolant injection that the coolant flow guiding capability in the cylindrical hole was apparently established after five hole diameters and no significant changes in the film cooling effectiveness distribution could be observed for the greater L/D. However, the film cooling effectiveness characteristics generally decreased with decreasing hole L/D ratio in the range of 1.75 ≤ L/D ≤ 5.0. This decrease in film cooling performance was attributed to (1) the undeveloped character of the flow in the coolant channels and (2) the greater effective injection angle of the coolant flow with respect to the external flow direction and surface. The lowest values of film cooling effectiveness were measured for the smallest hole length-to-diameter ratio, L/D = 1.75.

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Friedrichs, S., H. P. Hodson, and W. N. Dawes. "The Design of an Improved Endwall Film-Cooling Configuration." Journal of Turbomachinery 121, no. 4 (October 1, 1999): 772–80. http://dx.doi.org/10.1115/1.2836731.

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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, while 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 modeling developed by 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, while 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): First, coolant ejection downstream of the three-dimensional separation lines on the endwall does not change the secondary flow structures; second, 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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Gritsch, Michael, Achmed Schulz, and Sigmar Wittig. "Effect of Internal Coolant Crossflow on the Effectiveness of Shaped Film-Cooling Holes." Journal of Turbomachinery 125, no. 3 (July 1, 2003): 547–54. http://dx.doi.org/10.1115/1.1580523.

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Film-cooling was the subject of numerous studies during the past decades. However, the effect of flow conditions on the entry side of the film-cooling hole on film-cooling performance has surprisingly not received much attention. A stagnant plenum which is widely used in experimental and numerical studies to feed the holes is not necessarily a right means to re-present real engine conditions. For this reason, the present paper reports on an experimental study investigating the effect of a coolant crossflow feeding the holes that is oriented perpendicular to the hot gas flow direction to model a flow situation that is, for instance, of common use in modern turbine blades’ cooling schemes. A comprehensive set of experiments was performed to evaluate the effect of perpendicular coolant supply direction on film-cooling effectiveness over a wide range of blowing ratios (M=0.5…2.0) and coolant crossflow Mach numbers Mac=0…0.6. The coolant-to-hot gas density ratio, however, was kept constant at 1.85 which can be assumed to be representative for typical gas turbine applications. Three different hole geometries, including a cylindrical hole as well as two holes with expanded exits, were considered. Particularly, two-dimensional distributions of local film-cooling effectiveness acquired by means of an infrared camera system were used to give detailed insight into the governing flow phenomena. The results of the present investigation show that there is a profound effect of how the coolant is supplied to the hole on the film-cooling performance in the near hole region. Therefore, crossflow at the hole entry side has be taken into account when modeling film-cooling schemes of turbine bladings.

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Wang, J. H., J. Messner, and H. Stetter. "An Experimental Investigation on Transpiration Cooling Part II: Comparison of Cooling Methods and Media." International Journal of Rotating Machinery 10, no. 5 (2004): 355–63. http://dx.doi.org/10.1155/s1023621x04000363.

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This article attempts to provide a cooling performance comparison of various mass transfer cooling methods and different cooling media through two experiments. In the first experiment, pressurized air was used as a cooling medium and two different circular tubes were used as specimens. One is made of impermeable solid material with four rows of discrete holes to simulate film cooling, and the other consists of sintered porous material to create a porous transpiration cooling effect. The natures of transpiration cooling and film cooling including leading and trailing edge injection cooling were compared. This experiment found that by using a gaseous cooling medium, transpiration cooling could provide a higher cooling effect and a larger coolant coverage than film cooling in the leading stagnation region, and on the side of the specimen at the same coolant injection flow rates; but in the trailing stagnation region, the traditional coolant injection method through discrete film holes might be better than transpiration cooling, especially for turbine blades with thin trailing edges. In the second experiment, the cooling effects of gaseous and liquid media on the same porous tube's surface were compared. This experiment showed that the porous areas cooled using gaseous and liquid cooling media were almost identical, but the cooling effect of liquid evaporation was much higher than that of gaseous cooling, especially in the leading and trailing stagnation regions of turbine blades. This important discovery makes it possible to solve the stagnation region problems in turbine blade cooling.

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Bazdidi-Tehrani, F., and G. E. Andrews. "Full-Coverage Discrete Hole Film Cooling: Investigation of the Effect of Variable Density Ratio." Journal of Engineering for Gas Turbines and Power 116, no. 3 (July 1, 1994): 587–96. http://dx.doi.org/10.1115/1.2906860.

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Experimental results of the overall and adiabatic cooling effectiveness for full-coverage discrete hole film cooling are presented for a range of practical geometries. The results are reported for various hot gas mainstream-to-coolant temperature (density) ratios, in the realistic range of 1.0–3.2. The variation of this ratio was achieved by increasing the crossflow mainstream temperature, over the range 300–930 K. For combustor wall film cooling applications, the overall cooling effectiveness increased significantly with the number of holes per unit wall surface area, over the range of 4306–26910 m−2 and with the hole size, in the range of 1.0–2.2 mm, due to the improvement in film cooling. The effect of varying the mainstream-to-coolant temperature ratio, in the present range of 1.0–3.2, on the film cooling performance was shown to be small and no consistent trends were established for various configurations, for the coolant mass flow rates per unit wall surface area, less than 0.4 kg/sm2. At a higher value of 0.89 kg/sm2, an increase in the temperature ratio improved the film cooling performance slightly.

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Shangguan, Yanqin, and Fei Cao. "An LBM-Based Investigation on the Mixing Mechanism of Double Rows Film Cooling with the Combination of Forward and Backward Jets." Energies 15, no. 13 (July 1, 2022): 4848. http://dx.doi.org/10.3390/en15134848.

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Film cooling has been widely applied to the highly efficient thermal protection of gas turbines. By using the simplified thermal lattice Boltzmann method (STLBM), a series of large-scale simulations of film cooling are performed to dig up the mixing mechanism of double rows film cooling with the combination of forward and backward jets at the first attempt. The combination of an upstream row with forward jet and a downstream row with backward jet is considered. The Reynolds number is 4000. The blowing ratio of the upstream coolant jet is fixed as BR1=0.5. For the downstream coolant jet (BR2), five values ranging from 0.2–0.8 are considered. The inclination angles of forward jet and backward jet are 35° and 145°, respectively. The numerical results reveal that the performance of film cooling is greatly improved by backward downstream jet due to the suppression of counterrotating vortex pair (CVP). Moreover, the flow structure is changed with the blowing ratio of backward jet. An anti-CVP having the opposite rotational direction to CVP appears as the blowing ratio of backward jet is large. The special flow structure weakens the adverse effect of CVP and transports much coolant jet to the cooled wall. Correspondingly, the time-averaged film cooling effectiveness is increased and the fluctuation of film cooling effectiveness is decreased. All of these indicate that a backward downstream jet with a large blowing ratio improves film cooling performance. The results obtained in this work help to the optimization of film cooling scheme, which also benefit the promotion and application of STLBM in gas turbine engineering.

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Dissertations / Theses on the topic "Film cooling"

1

Hossain, Mohammad Arif. "Sweeping Jet Film Cooling." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1586462423029754.

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Strater, Kurt F. "Countercurrent cooling of blown film." Thesis, McGill University, 1985. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=66003.

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Fawcett, Richard James. "Coherent unsteadiness in film cooling." Thesis, University of Oxford, 2011. http://ora.ox.ac.uk/objects/uuid:57ec3da6-4946-4f66-8421-b01d53d7e0fc.

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Film cooling is vital for the cooling of the blades and vanes in the high temperature environment of a jet engine high pressure turbine stage. Previous research into film cooling has typically concentrated on its time-mean performance. However, results from other studies upon more simplified geometries, suggest that coherent unsteadiness is likely to also be present in film cooling flows. The research presented in this thesis, therefore, aims to characterise what coherent unsteadiness, if any, is present within film cooling flows. Cylindrical and shaped cooling holes, located upon the pressure surface of a turbine blade within a large scale linear cascade, have been investigated. A blowing ratio range of 0.5 to 2.0 has been investigated, with either a plenum or perpendicular crossflow at the cooling hole inlet. Particle Image Velocimetry, high speed photography and Hot Wire Anemometry have been used to investigate the jet downstream of both cooling holes. The impact of crossflow at the hole inlet upon the flowfield inside both cooling holes has been investigated using Hot Wire Anemometry and a further numerical model solved by applying TBLOCK. The results presented in the current thesis, show the existence of two coherent unsteady structures in the jet downstream of both the cylindrical and the shaped holes. These structures are called shear layer vortices and hairpin vortices, and their formation is dependent on the velocity difference across the jet shear layer. Inside the cooling hole coherent hairpin vortices also appear to occur, with their formation dependent on the direction and magnitude of the crossflow at the hole inlet. The coherent unsteadiness presented here is shown for the first time for film cooling flows, and recommendations to build on the current study, in what is potentially an interesting research area, are made at the end of this thesis.

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Aghasi, Paul P. "Dependence of Film Cooling Effectiveness on 3D Printed Cooling Holes." University of Cincinnati / OhioLINK, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1458893416.

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Nowlin, Scott Raymond. "The use of intersecting film cooling passages for nozzle guide vane cooling." Thesis, University of Oxford, 2009. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.670018.

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Licu, Dragos N. "Heat transfer characteristics in film cooling applications." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1998. http://www.collectionscanada.ca/obj/s4/f2/dsk2/tape17/PQDD_0005/NQ34581.pdf.

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Friedrichs, Stefan. "Endwall film-cooling in axial flow turbines." Thesis, University of Cambridge, 1997. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.627225.

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Telisinghe, Janendra C. "Film cooling of turbine blade trailing edges." Thesis, University of Oxford, 2013. http://ora.ox.ac.uk/objects/uuid:86c06246-16e9-4378-9a61-e09317d31a92.

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In modern gas turbine engines, film cooling is extensively used to cool the components exposed to the hot mainstream gas path. In implementing film cooling on modern gas turbine engines, the trailing edge film poses a particularly challenging design problem. From an aerodynamic point of view, the trailing edge of a blade is designed to be as thin as possible. However, this conflicts with the implementation of the cooling design. The most common method of film cooling the trailing edge is via late pressure surface discrete film cooling holes. Another method of cooling the trailing edge is by using discrete pressure surface slots. This thesis documents a comparative aerodynamic and heat transfer study of three trailing edge cooling configurations. The study was carried out using a large scale, low speed wind tunnel situated at the Southwell Laboratory. The three trailing edge cooling configurations considered were as follows. First is the common late pressure film cooling of the trailing edge via discrete film cooling holes. This configuration is designated as datum configuration. Second is the pressure surface slot coolant ejection. This configuration was designated as cast cutback configuration. The third is the pressure surface ejection via discrete film cooling holes within a step cutback. This configuration was designated the machined cutback configuration. The above configurations were incorporated into three flat plates manufactured using stereolithography. In the aerodynamic study, the static pressure distribution and discharge coefficient for the three configurations were compared. Furthermore, two dimensional total pressure measurements were carried out using a traverse mechanism downstream of the test plates. The total pressure measurements were used to compute the mixed out losses for the three configurations. It was found that the datum and machined cutback configurations have similar discharge coefficients and mixed out losses whilst the cast cutback configuration produces greater mixed out loss. The film effectiveness and heat transfer coefficient on the pressure surface downstream of the coolant ejection was obtained using a steady state heat transfer technique. The effectiveness measurements were compared with those from the literature and correlated against the two dimensional slot model. The heat transfer measurements show that the cast cutback configuration has the potential to give higher effectiveness at the trailing edge.

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Renze, Peter Clemens-August. "Large eddy simulation of film cooling flows /." Aachen : Shaker, 2008. http://d-nb.info/989959031/04.

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Kheniser, Issam E. "Film Cooling Experiments in a Medium Duration Blowdown Facility." The Ohio State University, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=osu1276540410.

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Books on the topic "Film cooling"

1

Aerssens, S. E. E. Modelling of two-dimensional film cooling. Manchester: UMIST, 1994.

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Simon, Frederick F. Jet model for slot film cooling with effect of free-stream and coolant turbulence. Cleveland, Ohio: Lewis Research Center, 1986.

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Nan-Suey, Liu, and NASA Glenn Research Center, eds. Film cooling flow effects on post-combustor trace chemistry. Cleveland, Ohio: National Aeronautics and Space Administration, Glenn Research Center, 2003.

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4

Garg, Vijay Kumar. Effect of coolant temperature and mass flow on film cooling of turbine blades. [Washington, D.C: National Aeronautics and Space Administration, 1997.

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Olsen, George C. Hydrogen film cooling with incident and swept-shock interactions in a Mach 6.4 nitrogen free stream. Hampton, Va: Langley Research Center, 1995.

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Acharya, Sumanta. Large eddy simulations and turbulence modeling for film cooling. [Cleveland, Ohio]: National Aeronautics and Space Administration, Glenn Research Center, 1999.

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Ward, S. C. Validation of a CFD model for predicting film cooling performance. Washington, D. C: American Institute of Aeronautics and Astronautics, 1993.

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Garg, Vijay Kumar. Leading edge film cooling effects on turbine blade heat transfer. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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9

J, Bruckner R., Smith F. A, and United States. National Aeronautics and Space Administration., eds. Application of thin-film thermocouples to localized heat transfer measurements. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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Lepicovsky, J. Application of thin-film thermocouples to localized heat transfer measurements. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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Book chapters on the topic "Film cooling"

1

Bogard, David G., and Karen A. Thole. "Film Cooling." In Turbine Aerodynamics, Heat Transfer, Materials, and Mechanics, 223–73. Reston, VA: American Institute of Aeronautics and Astronautics, Inc., 2014. http://dx.doi.org/10.2514/5.9781624102660.0223.0274.

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Herwig, Heinz. "Filmkühlung (film cooling)." In Wärmeübertragung A-Z, 62–66. Berlin, Heidelberg: Springer Berlin Heidelberg, 2000. http://dx.doi.org/10.1007/978-3-642-56940-1_14.

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Peter, Johannes M. F., and Markus J. Kloker. "Numerical Simulation of Film Cooling in Supersonic Flow." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 79–95. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_5.

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Abstract High-order direct numerical simulations of film cooling by tangentially blowing cool helium at supersonic speeds into a hot turbulent boundary-layer flow of steam (gaseous H2O) at a free stream Mach number of 3.3 are presented. The stagnation temperature of the hot gas is much larger than that of the coolant flow, which is injected from a vertical slot of height s in a backward-facing step. The influence of the coolant mass flow rate is investigated by varying the blowing ratio F or the injection height s at kept cooling-gas temperature and Mach number. A variation of the coolant Mach number shows no significant influence. In the canonical baseline cases all walls are treated as adiabatic, and the investigation of a strongly cooled wall up to the blowing position, resembling regenerative wall cooling present in a rocket engine, shows a strong influence on the flow field. No significant influence of the lip thickness on the cooling performance is found. Cooling correlations are examined, and a cooling-effectiveness comparison between tangential and wall-normal blowing is performed.

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Ludescher, Sandra, and Herbert Olivier. "Film Cooling in Rocket Nozzles." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 65–78. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_4.

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Abstract In this project supersonic, tangential film cooling in the expansion part of a nozzle with rocket-engine like hot gas conditions was investigated. Therefore, a parametric study in a conical nozzle was conducted revealing the most important influencing parameter on film cooling for the presented setup. Additionally, a new axisymmetric film cooling model and a method for calculating the cooling efficiency from experimental data was developed. These models lead to a satisfying correlation of the data. Furthermore, film cooling in a dual-bell nozzle performing in altitude mode was investigated. The aim of these experiments was to show the influence of different contour inflection geometries on the film cooling efficiency in the bell extension.

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Guo, Tingting, Shaohua Li, and Jianhong Liu. "Large Eddy Simulation of Film Cooling." In Challenges of Power Engineering and Environment, 1419–22. Berlin, Heidelberg: Springer Berlin Heidelberg, 2007. http://dx.doi.org/10.1007/978-3-540-76694-0_267.

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Suresh, Batchu, Resham D. Khade, V. Kesavan, and D. Kishore Prasad. "Investigation for the Improvement of Film Cooling Effectiveness of Effusion Cooling Holes." In Proceedings of the National Aerospace Propulsion Conference, 171–84. Singapore: Springer Singapore, 2020. http://dx.doi.org/10.1007/978-981-15-5039-3_10.

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Kulju, Timo, Juha Pyykkönen, David C. Martin, Esa Muurinen, and Riitta L. Keiski. "CFD-Simulation of Film Boiling at Steel Cooling Process." In Film and Nucleate Boiling Processes, 28–44. 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959: ASTM International, 2011. http://dx.doi.org/10.1520/stp49331t.

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Kulju, Timo, Juha Pyykkönen, David C. Martin, Esa Muurinen, and Riitta L. Keiski. "CFD-Simulation of Film Boiling at Steel Cooling Process." In Film and Nucleate Boiling Processes, 28–44. 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959: ASTM International, 2011. http://dx.doi.org/10.1520/stp153420120002.

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Medic, Gorazd, and Paul Durbin. "RANS Simulations for Film-Cooling Analysis and Design." In Modelling Fluid Flow, 213–28. Berlin, Heidelberg: Springer Berlin Heidelberg, 2004. http://dx.doi.org/10.1007/978-3-662-08797-8_15.

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Zhang, Ling, Tingting Guo, Shaohua Li, and Jianhong Liu. "Effects of Jet Geometries on Film-cooling Effectiveness." In Challenges of Power Engineering and Environment, 1357–60. Berlin, Heidelberg: Springer Berlin Heidelberg, 2007. http://dx.doi.org/10.1007/978-3-540-76694-0_255.

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Conference papers on the topic "Film cooling"

1

Takeishi, K., Y. Oda, Y. Egawa, and T. Kitamura. "Film cooling with swirling coolant flow." In HEAT TRANSFER 2010. Southampton, UK: WIT Press, 2010. http://dx.doi.org/10.2495/ht100171.

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BASS, ROBERT, LARRY HARDIN, RICHARD RODGERS, and RICHARD ERNST. "Supersonic film cooling." In 2nd International Aerospace Planes Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1990. http://dx.doi.org/10.2514/6.1990-5239.

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3

Yusop, Nadiahnor Md, Gordon E. Andrews, Derek B. Ingham, I. M. Khalifa, Mike C. Mkpadi, and Mohammed Pourkashanian. "Predictions of Adiabatic Film Cooling Effectiveness for Effusion Film Cooling." In ASME Turbo Expo 2007: Power for Land, Sea, and Air. ASMEDC, 2007. http://dx.doi.org/10.1115/gt2007-27467.

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This paper presents computational predictions of adiabatic film cooling effectiveness for effusion cooling systems with 90° and 30° holes. Predictions are performed for a range of coolant injection mass flow rates per unit surface area, G, of 0.1kg/sm2 - 1.6 kg/sm2 for 90° holes with constant pitch-to-diameter ratio of X/D = 11 and 10 rows of holes and for 30° inclined holes with X/D = 11 and 15 rows of holes over a 152mm surface length. The computational works performed are steady-state and the turbulent governing equations are solved by a control-volume-based finite difference method with second-order upwind scheme and the k-epsilon turbulence model. The velocity and pressure terms of momentum equations are solved by the SIMPLE method. The CFD prediction were validated by comparing the predictions with literature data for single rows of inclined holes and then applied to effusion cooling. The predictions included the use of a tracer gas in the coolant, which was used to predict the mixing of the coolant with the hot mainstream gases. Also the surface distribution of the tracer gas was a direct prediction of the cooling effectiveness. The mixing of coolant with the mainstream was studied and boundary layer temperature and coolant mixing profiles were predicted. These were compared with temperature measurement in a hot effusion cooling test rig.

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4

Bai, Jiang-Tao, Hui-ren Zhu, and Cun-liang Liu. "Film Cooling Characteristic of Double-Fan-Shaped Film Cooling Holes." In ASME Turbo Expo 2009: Power for Land, Sea, and Air. ASMEDC, 2009. http://dx.doi.org/10.1115/gt2009-59318.

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The film cooling performance downstream of a single row of double-fan-shaped film cooling holes in a flat plate have been investigated by experimental measurements and numerical simulation. The entrance and exit of double-fan-shaped holes are comprised of a lateral expansion of 15° from the original simple cylindrical shape with stream-wise inclination of 45°. The width of the exit face to cylinder diameter ratio is 1.5; the length-to-diameter ratio is 4.24 and the pitch-to-diameter ratio is 3. The experimental method used to obtain the adiabatic film cooling effectiveness values and the heat transfer coefficient is a transient narrow band liquid crystal technique. Both film cooling effectiveness and heat transfer coefficient are measured at three momentum ratios (I = 0.5, 1, 2) at constant Reynolds number (Re = 10000) and free stream turbulence (Tu = 2%). The film cooling effectiveness, heat transfer coefficient and Net Heat Flux Reduction (NHFR) are presented for detailed distribution and span-wise averaged values. Discharge coefficients are also measured in the experiment. A commercial package is used to numerically simulate the flow and heat transfer of double-fan-shaped holes; simple cylindrical holes are also simulated for comparison. Numerical simulation use RNG turbulence model with a standard wall function for near wall region. Experimental and Numerical simulation results show that: 1) the double-fan-shaped holes present higher discharge coefficient than simple cylindrical holes at respective momentum ratio; 2) the numerical simulation film cooling effectiveness results of double-fan-shaped holes accord well with the experimental results; 3) at measured three momentum ratios, the double-fan-shaped holes demonstrate better film cooling performance (higher NHFR) than simple cylindrical holes, better film cooling expansion on span-wise; 4) the best momentum ratio of double-fan-shaped holes is 0.5.

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Natsui, Greg, Zachary Little, Jay Kapat, Anthony Socotch, Anquan Wang, and Jason E. Dees. "Adiabatic Film Cooling Effectiveness Measurements Throughout Multi-Row Film Cooling Arrays." In ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2016. http://dx.doi.org/10.1115/gt2016-56183.

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Adiabatic film cooling effectiveness measurements are obtained using pressure-sensitive paint (PSP) on a flat film cooled surface. The effects of blowing ratio and hole spacing are investigated for four multi-row arrays comprised of 8 rows containing 52 holes of 3.8 mm diameter with 20° inclination angles and hole length-to-diameter ratio of 11.2. The four arrays investigated have two different hole-to-hole spacings composed of cylindrical and diffuser holes. For the first case, lateral and streamwise pitches are 7.5 times the diameter. For the second case, pitch-to-diameter ratio is 14 in lateral direction and 10 in the streamwise direction. The holes are in a staggered arrangement. Adiabatic effectiveness measurements are taken for a blowing ratio range of 0.3 to 1.2 and a density ratio of 1.5, with CO2 injected as the coolant. A thorough boundary layer analysis is presented, and data was taken using hotwire anemometry with air injection, with boundary layer and turbulence measurements taken at multiple locations in order to characterize the boundary layer. Local effectiveness, laterally averaged effectiveness, boundary layer thickness, momentum thickness, turbulence intensity and turbulence length scale are presented. For the cylindrical holes, at the first row of injection, the film jets are still attached at a blowing ratio of 0.3. By a blowing ratio of 0.5, the jet is observed to lift off, and then impinge back onto the test surface. At a blowing ratio of 1.2, the jets lift off, but reattach much further downstream, spreading the coolant further along the test surface. A thorough uncertainty analysis has been conducted in order to fully understand the presented measurements and any shortcomings of the measurement technique. The maximum uncertainty of effectiveness and blowing ratio is 0.02 counts of effectiveness and 3 percent respectively.

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6

Rhee, Dong Ho, Youn Seok Lee, and Hyung Hee Cho. "Film Cooling Effectiveness and Heat Transfer of Rectangular-Shaped Film Cooling Holes." In ASME Turbo Expo 2002: Power for Land, Sea, and Air. ASMEDC, 2002. http://dx.doi.org/10.1115/gt2002-30168.

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An experimental study has been conducted to measure the local film-cooling effectiveness and the heat transfer coefficient for a single row of rectangular-shaped holes. The holes have a 35° inclination angle with 3 hole diameter spacing of rectangular cross-sections. Four different cooling hole shapes such as a straight rectangular hole, a rectangular hole with laterally expanded exit, a circular hole and a two-dimensional slot are tested. The rectangular cross-section has the aspect ratio of 2 at the hole inlet with the hydraulic diameter of 10 mm. The area ratio of the exit to the hole inlet is 1.8 for the rectangular hole with expanded exit, which is similar to a two-dimensional slot. A thermochromic liquid crystals technique is applied to determine adiabatic film cooling effectiveness values and heat transfer coefficients on the test surface. Both film cooling effectiveness and heat transfer coefficient are measured for various blowing rates and compared with the results of the cylindrical holes and the two-dimensional slot. The flow patterns inside and downstream of holes are calculated numerically by a commercial package. The results show that the rectangular holes provide better performance than the cylindrical holes. For the rectangular holes with laterally expanded exit, the penetration of jet is reduced significantly, and the higher and more uniform cooling performance is obtained even at relatively high blowing rates. The reason is that the rectangular hole with expanded exit reduces momentum of coolant and promotes the lateral spreading like a two-dimensional slot.

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7

Kim, Sun-min, Ki-Don Lee, and Kwang-Yong Kim. "Numerical Study on Film-Cooling Effectiveness for Various Film-Cooling Hole Schemes." In ASME-JSME-KSME 2011 Joint Fluids Engineering Conference. ASMEDC, 2011. http://dx.doi.org/10.1115/ajk2011-22003.

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Film-cooling has been widely used as the important alternative to protect the turbine blade. Since the film-cooling hole geometry is one of the most influential parameters for film-cooling performance, various film-cooling hole schemes have been developed to increase cooling performance for the past few decades. In the present work, numerical analysis has been performed to investigate and to compare the film-cooling performance of various film-cooling hole schemes such as fan-shaped, crescent, louver, and dumbbell holes. For analyzes of the turbulent flow and film-cooling, three-dimensional Reynolds-averaged Navier-Stokes analysis has been performed with shear stress transport turbulence model. The validation of numerical results has been performed in comparison with experimental data. The flow characteristics and film-cooling performance for each hole shape have been investigated and evaluated in terms of local- and averaged film-cooling effectivenesses.

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Takeishi, Kenichiro, Yutaka Oda, Yuta Egawa, and Satoshi Hada. "Film Cooling With Swirling Coolant Flow Controlled by Impingement Cooling in a Closed Cavity." In ASME 2011 Power Conference collocated with JSME ICOPE 2011. ASMEDC, 2011. http://dx.doi.org/10.1115/power2011-55390.

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A new film cooling concept has been developed by managing the swirled film coolant induced inside a hexagonal plenum by two slant impingement jets, which are inclined at α degree toward the vertical direction and installed in a staggered position on the plenum chamber wall. Film cooling tests have been conducted by using a circular film cooling hole model mounted on a low speed wind tunnel. Heat transfer coefficient distributions of inclined jet impingements in a closed cavity was measured by naphthalene sublimation method and the film cooling effectiveness on the surface of the wind tunnel was measured by pressure sensitive paint (PSP). It appeared from experimental results that the swirled film coolant flow deteriorated the film cooling effectiveness at low swirl number but improved it at high swirl number. To investigate the mechanism of the improved film cooling effectiveness by the swirled coolant, the spatial distribution of the film cooling effectiveness and flow field were measured by laser induced fluorescence (LIF) and particle image velocimetry (PIV), respectively. The coolant jet penetration into mainstream is suppressed by the strong swirling motion of the coolant. As a result the film cooling effectiveness distribution on the wall keeps higher value behind the cooling hole over a long range. Additionally, kidney vortex structure was disappeared at high swirl number.

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9

Holgate, Nicholas E., Peter T. Ireland, and Eduardo Romero. "The Effects of Combustor Cooling Features on Nozzle Guide Vane Film Cooling Experiments." In ASME Turbo Expo 2018: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2018. http://dx.doi.org/10.1115/gt2018-75249.

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Recent advances in experimental methods have allowed researchers to study nozzle guide vane film cooling in the presence of combustor dilution ports and endwall films. The dilution injection creates nonuniformities in temperature, velocity, and turbulence, and an understanding of the vane film cooling performance is complicated by competing influences. In this study, dilution port temperature profiles have been measured in the absence of vane film cooling and compared to film effectiveness measurements in the presence of both films and dilution, illustrating the effects of the dilution port turbulence on film cooling performance. It is found that dilution port injection can create significant effectiveness benefits at the difficult-to-cool vane stagnation region, due to the more turbulent hot mainstream enhancing the mixing of film coolant jets that have left the airfoil surface. Also explored are the implications of endwall film cooling for infrared vane surface temperature measurements. The reduced endwall temperatures reduce the thermal emissions from this surface, so reducing the amount of extraneous radiation reflected from the vane surface where measurements are being made. The results of a detailed calibration show that the maximum local film effectiveness measurement error could be up to 0.05 if this effect were to go unaccounted for.

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10

Hale, C. A., M. W. Plesniak, and S. Ramadhyani. "Film Cooling Effectiveness for Short Film Cooling Holes Fed by a Narrow Plenum." In ASME 1999 International Gas Turbine and Aeroengine Congress and Exhibition. American Society of Mechanical Engineers, 1999. http://dx.doi.org/10.1115/99-gt-036.

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The adiabatic, steady-state liquid crystal technique was used to measure surface adiabatic film cooling effectiveness values in the near-hole region (X / D < 10). A parametric study was conducted for a single row of short holes (L / D ≤ 3) fed by a narrow plenum (H / D = 1). Film cooling effectiveness values are presented and compared for various L / D ratios (0.66 to 3.0), three different blowing ratios (0.5, 1.0, and 1.5), two different plenum feed configurations (co-flow and counter flow), and two different injection angles (35° and 90°). Injection hole geometery and plenum feed direction were found to significantly affect short hole film cooling performance. Under certain conditions, comparable or improved coverage was achieved with 90° holes as with 35° holes. This result has important implications for manufacturing of thin-walled film-cooled blades or vanes.

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Reports on the topic "Film cooling"

1

Wygnanski, Israel J., Alfonso Ortega, and Hermann Fasel. Film Cooling by a Pulsating Wall Jet. Fort Belvoir, VA: Defense Technical Information Center, June 1997. http://dx.doi.org/10.21236/ada329651.

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2

Leylek, James H., D. K. Walters, William D. York, D. S. Holloway, and Jeffrey D. Ferguson. Computational Film Cooling Methods for Gas Turbine Airfoils. Fort Belvoir, VA: Defense Technical Information Center, March 2002. http://dx.doi.org/10.21236/ada400186.

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3

Muss, Jeffrey. 5Klbf Unielement TCA for Film Cooling Model Validation. Fort Belvoir, VA: Defense Technical Information Center, April 2003. http://dx.doi.org/10.21236/ada414443.

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Himansu, Ananda, Edward B. Coy, Venkateswaran Sankaran, and Steven A. Danczyk. Modeling of Fuel Film Cooling on Chamber Hot Wall. Fort Belvoir, VA: Defense Technical Information Center, July 2014. http://dx.doi.org/10.21236/ada611830.

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Schauer, John J., and David J. Pestian. Film Cooling Heat Transfer with High Free Stream Turbulence. Fort Belvoir, VA: Defense Technical Information Center, November 1995. http://dx.doi.org/10.21236/ada312458.

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Ramesh, Sridharan, and Douglas Straub. Film Cooling Performance Predictions For Air And Supercritical CO2. Office of Scientific and Technical Information (OSTI), October 2021. http://dx.doi.org/10.2172/1827885.

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Coulthard, Sarah M. Effects of Pulsing on Film Cooling of Gas Turbine Airfoils. Fort Belvoir, VA: Defense Technical Information Center, May 2005. http://dx.doi.org/10.21236/ada437128.

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Raben, Sam, Pavlos Vlachos, and Wing Ng. Effects of Leading Edge Film-Cooling and Surface Roughness on the Downstream Film-Cooling Along a Transonic Turbine Blade for Low and High Free-Stream Turbulence. Fort Belvoir, VA: Defense Technical Information Center, February 2008. http://dx.doi.org/10.21236/ada479415.

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9

Bogard, David G., and Karen A. Thole. Improving Durability of Turbine Components Through Trenched Film Cooling and Contoured Endwalls. Office of Scientific and Technical Information (OSTI), September 2014. http://dx.doi.org/10.2172/1224799.

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10

Wang, L., H. Tsang, T. Simon, and E. Eckert. Measurements of mean flow and eddy transport over a film cooling surface. Office of Scientific and Technical Information (OSTI), May 1996. http://dx.doi.org/10.2172/251410.

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