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Relevant bibliographies by topics / Heat-transfer; Aerospace gas turbines

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Academic literature on the topic 'Heat-transfer; Aerospace gas turbines'

Author: Grafiati

Published: 4 June 2021

Last updated: 10 February 2022

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Journal articles on the topic "Heat-transfer; Aerospace gas turbines"

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Fersini, Maurizio, R. Bianco, L. De Lorenzis, Antonio Licciulli, G. Pasquero, and G. Zanon. "Thermo-Structural Analysis of Ceramic Vanes for Gas Turbines." Advances in Science and Technology 45 (October 2006): 1759–64. http://dx.doi.org/10.4028/www.scientific.net/ast.45.1759.

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Advanced structural ceramics such as Hot Pressed Silicon Nitride (HPSN) and Reaction Bonded Silicon Carbide (RBSC), thanks to their low density (3.1 ÷ 3.4 gr/cm3) and to their thermostructural properties, are interesting candidates for aerospace applications. This research investigates the feasibility of employing such monolithic advanced ceramics for the production of turbine vanes for aerospace applications, by means of a finite element analysis. A parametric study is performed to analyse the influence of the coefficient of thermal expansion, the specific heat, the thermal conductivity, and the Weibull modulus on structural stability, heat transfer properties and thermomechanical stresses under take-off and flying conditions. A nodal point that is evidenced is the high intensity of thermal stresses on the vane, both on steady state and in transient conditions. In order to reduce such stresses various simulations have been carried out varying geometrical parameters such as the wall thickness. Several open questions are evidenced and guidelines are drawn for the design and production of ceramic vanes for gas turbines.

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Lock, Gary D., Michael Wilson, and J. Michael Owen. "Influence of Fluid Dynamics on Heat Transfer in a Preswirl Rotating-Disk System." Journal of Engineering for Gas Turbines and Power 127, no. 4 (March 1, 2004): 791–97. http://dx.doi.org/10.1115/1.1924721.

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Modern gas turbines are cooled using air diverted from the compressor. In a “direct-transfer” preswirl system, this cooling air flows axially across the wheel space from stationary preswirl nozzles to receiver holes located in the rotating turbine disk. The distribution of the local Nusselt number Nu on the rotating disk is governed by three nondimensional fluid-dynamic parameters: preswirl ratio βp, rotational Reynolds number Reϕ, and turbulent flow parameter λT. This paper describes heat transfer measurements obtained from a scaled model of a gas turbine rotor-stator cavity, where the flow structure is representative of that found in the engine. The experiments reveal that Nu on the rotating disk is axisymmetric except in the region of the receiver holes, where significant two-dimensional variations have been measured. At the higher coolant flow rates studied, there is a peak in heat transfer at the radius of the preswirl nozzles associated with the impinging jets from the preswirl nozzles. At lower coolant flow rates, the heat transfer is dominated by viscous effects. The Nusselt number is observed to increase as either Reϕ or λT increases.

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Valenti, Michael. "Turbines for Peace." Mechanical Engineering 122, no. 08 (August 1, 2000): 70–72. http://dx.doi.org/10.1115/1.2000-aug-5.

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This article discusses that military-sponsored research tools can improve the machines that drive civil applications. The Defense Evaluation and Research Agency (DERA) researchers tested the engine of the legendary DeHavilland Vampire single seat jet fighter in the late 1940s. This Vampire is owned by Fred Ihlenburg, president of Yakity Yaks Inc., an importer of foreign military aircraft, based in Aurora, Oregon. DERA is investigating heat transfer on turbine blades to help gas turbine manufacturers develop a cooling system that will keep blades at an optimum temperature while minimizing losses in engine performance. More efficient cooling means less air is bled from the compressor, thus improving performance while extending blade life. This work was co-funded by the Central European Commission under the Brite Euram Fourth-Framework Initiative, which is part of the European Union’s strategy to enhance European global competitiveness, and Britain’s Department of Trade and Industry’s Civil Aircraft Research and Technology Demonstration Program. The British program aims to advance the capabilities of the United Kingdom’s civil aerospace companies.

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Rice, I. G. "Thermodynamic Evaluation of Gas Turbine Cogeneration Cycles: Part I—Heat Balance Method Analysis." Journal of Engineering for Gas Turbines and Power 109, no. 1 (January 1, 1987): 1–7. http://dx.doi.org/10.1115/1.3240001.

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This paper presents a heat balance method of evaluating various open-cycle gas turbines and heat recovery systems based on the first law of thermodynamics. A useful graphic solution is presented that can be readily applied to various gas turbine cogeneration configurations. An analysis of seven commercially available gas turbines is made showing the effect of pressure ratio, exhaust temperature, intercooling, regeneration, and turbine rotor inlet temperature in regard to power output, heat recovery, and overall cycle efficiency. The method presented can be readily programmed in a computer, for any given gaseous or liquid fuel, to yield accurate evaluations. An X–Y plotter can be utilized to present the results.

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Kurowski, Marcin, Ryszard Szwaba, Janusz Telega, Pawel Flaszynski, Fernando Tejero, and Piotr Doerffer. "Wall distance effect on heat transfer at high flow velocity." Aircraft Engineering and Aerospace Technology 91, no. 9 (October 7, 2019): 1180–86. http://dx.doi.org/10.1108/aeat-01-2018-0022.

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Purpose This paper aims to present the results of experimental and numerical research on heat transfer distribution under the impinging jets at various distances from the wall and high jet velocity. This work is a part of the INNOLOT Program financed by National Centre for Research and Development. Design/methodology/approach The air jets flow out from the common pipe and impinge on a surface which is cooled by them, and in this way, all together create a model of external cooling system of low-pressure gas turbine casing. Measurements were carried out for the arrangement of 26 in-line jets with orifice diameter of 0.9 mm. Heat transfer distribution was investigated for various Reynolds and Mach numbers. The cooled wall, made of transparent PMMA, was covered with a heater foil on which a layer of self-adhesive liquid crystal foil was placed. The jet-to-wall distance was set to h = from 4.5 to 6 d. Findings The influence of various Reynolds and Mach numbers on cooled flat plate and jet-to-wall distance in terms of heat transfer effectiveness is presented. Experimental results used for the computational fluid dynamics (CFD) model development, validation and comparison with numerical results are presented. Practical implications Impinging air jets is a commonly used technique to cool advanced turbines elements, as it produces large convection enhancing the local heat transfer, which is a critical issue in the development of aircraft engines. Originality/value The achieved results present experimental investigations carried out to study the heat transfer distribution between the orthogonally impinging jets from long round pipe and flat plate. Reynolds number based on the jet orifice exit conditions was varied between 2,500 and 4,000; meanwhile, for such Re, the flow velocity in jets was particularly very high, changing from M = 0.56 to M = 0.77. Such flow conditions combination, i.e. the low Reynolds number and very high flow velocity cannot be found in the existing literature.

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Du, Haifen, Daimei Xie, Wei Jiang, Tong Chen, and Jianshu Gao. "Numerical Study on Heat Transfer Enhancement of Swirl Chamber on Gas Turbine Blade." International Journal of Turbo & Jet-Engines 35, no. 4 (December 19, 2018): 403–12. http://dx.doi.org/10.1515/tjj-2016-0049.

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Abstract The optimization of turbine cooling design has become a new research field of gas turbine. The swirl chamber is a prospect cooling concept. In this paper, the numerical simulation of the swirl chamber is carried out by FLUENT. The influence of inlet size parameters, temperature ratio and inlet Reynolds number on the enhanced heat transfer of swirl chamber is studied. The results show that, in the range of the studied condition, Nusselt number decreases with the height, the width, the ratio of width to height and Reynolds number. It also shows that comprehensive heat transfer effect is best at d=20 mm and enhances observably with the enlargement of width, width height ratio, and Reynolds number. Friction factor increases with height, width, temperature ratio and Reynolds number decreases. It is increased by increasing width height ratio. Nusselt number and comprehensive heat transfer effect decrease a little with aggrandizement of temperature ratio.

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Schiele, Ralf, and Sigmar Wittig. "Gas Turbine Heat Transfer: Past and Future Challenges." Journal of Propulsion and Power 16, no. 4 (July 2000): 583–89. http://dx.doi.org/10.2514/2.5611.

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Freedman, B. Z., and N. Lior. "A Novel High-Temperature Ejector-Topping Power Cycle." Journal of Engineering for Gas Turbines and Power 116, no. 1 (January 1, 1994): 1–7. http://dx.doi.org/10.1115/1.2906793.

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A novel, patented topping power cycle is described that takes its energy from a very high-temperature heat source and in which the temperature of the heat sink is still high enough to operate another, conventional power cycle. The top temperature heat source is used to evaporate a low saturation pressure liquid, which serves as the driving fluid for compressing the secondary fluid in an ejector. Due to the inherently simple construction of ejectors, they are well suited for operation at temperatures higher than those that can be used with gas turbines. The gases exiting from the ejector transfer heat to the lower temperature cycle, and are separated by condensing the primary fluid. The secondary gas is then used to drive a turbine. For a system using sodium as the primary fluid and helium as the secondary fluid, and using a bottoming Rankine steam cycle, the overall thermal efficiency can be at least 11 percent better than that of conventional steam Rankine cycles.

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Lock, Gary D., Youyou Yan, Paul J. Newton, Michael Wilson, and J. Michael Owen. "Heat Transfer Measurements Using Liquid Crystals in a Preswirl Rotating-Disk System." Journal of Engineering for Gas Turbines and Power 127, no. 2 (April 1, 2005): 375–82. http://dx.doi.org/10.1115/1.1787509.

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Preswirl nozzles are often used in gas turbines to deliver the cooling air to the turbine blades through receiver holes in a rotating disk. The distribution of the local Nusselt number, Nu, on the rotating disk is governed by three nondimensional fluid-dynamic parameters: preswirl ratio, βp, rotational Reynolds number, Reϕ, and turbulent flow parameter, λT. A scaled model of a gas turbine rotor–stator cavity, based on the geometry of current engine designs, has been used to create appropriate flow conditions. This paper describes how a thermochromic liquid crystal, in conjunction with a stroboscopic light and digital camera, is used in a transient experiment to obtain contour maps of Nu on the rotating disk. The thermal boundary conditions for the transient technique are such that an exponential-series solution to Fourier’s one-dimensional conduction equation is necessary. A method to assess the uncertainty in the measurements is discussed and these uncertainties are quantified. The experiments reveal that Nu on the rotating disk is axisymmetric except in the region of the receiver holes, where significant two-dimensional variations have been measured. At the higher coolant flow rates studied, there is a peak in heat transfer at the radius of the preswirl nozzles. The heat transfer is governed by two flow regimes: one dominated by inertial effects associated with the impinging jets from the preswirl nozzles, and another dominated by viscous effects at lower flow rates. The Nusselt number is observed to increase as either Reϕ or λT increases.

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Becker, B., and B. Schetter. "Gas Turbines Above 150 MW for Integrated Coal Gasification Combined Cycles (IGCC)." Journal of Engineering for Gas Turbines and Power 114, no. 4 (October 1, 1992): 660–64. http://dx.doi.org/10.1115/1.2906639.

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Commercial IGCC power plants need gas turbines with high efficiency and high power output in order to reduce specific installation costs and fuel consumption. Therefore the well-proven 154 MW V94.2 and the new 211 MW V94.3 high-temperature gas turbines are well suited for this kind of application. A high degree of integration of the gas turbine, steam turbine, oxygen production, gasifier, and raw gas heat recovery improves the cycle efficiency. The air use for oxygen production is taken from the gas turbine compressor. The N2 fraction is recompressed and mixed with the cleaned gas prior to combustion. Both features require modifications of the gas turbine casing and the burners. Newly designed burners using the coal gas with its very low heating value and a mixture of natural gas and steam as a second fuel are developed for low NOx and CO emissions. These special design features are described and burner test results presented.

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Dissertations / Theses on the topic "Heat-transfer; Aerospace gas turbines"

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Fletcher, Daniel Alden. "Internal cooling of turbine blades : the matrix cooling method." Thesis, University of Oxford, 1997. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.360259.

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Krumanaker, Matthew Lee. "Aerodynamics and Heat Transfer for a Modern Stage and One-Half Turbine." The Ohio State University, 2003. http://rave.ohiolink.edu/etdc/view?acc_num=osu1039538775.

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Kulkarni, Aditya Narayan. "Computational and Experimental Investigation of Internal Cooling Passages for Gas Turbine Applications." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1590591363859471.

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Lawson, Hannah. "Development of an Infrared Thermography Technique for Measuring Heat Transfer to a Flat Plate in a Blowdown Facility." The Ohio State University, 2015. http://rave.ohiolink.edu/etdc/view?acc_num=osu1429721463.

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Reding, Brian D. II. "Tubular and Sector Heat Pipes with Interconnected Branches for Gas Turbine and/or Compressor Cooling." FIU Digital Commons, 2013. http://digitalcommons.fiu.edu/etd/969.

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Designing turbines for either aerospace or power production is a daunting task for any heat transfer scientist or engineer. Turbine designers are continuously pursuing better ways to convert the stored chemical energy in the fuel into useful work with maximum efficiency. Based on thermodynamic principles, one way to improve thermal efficiency is to increase the turbine inlet pressure and temperature. Generally, the inlet temperature may exceed the capabilities of standard materials for safe and long-life operation of the turbine. Next generation propulsion systems, whether for new supersonic transport or for improving existing aviation transport, will require more aggressive cooling system for many hot-gas-path components of the turbine. Heat pipe technology offers a possible cooling technique for the structures exposed to the high heat fluxes. Hence, the objective of this dissertation is to develop new radially rotating heat pipe systems that integrate multiple rotating miniature heat pipes with a common reservoir for a more effective and practical solution to turbine or compressor cooling. In this dissertation, two radially rotating miniature heat pipes and two sector heat pipes are analyzed and studied by utilizing suitable fluid flow and heat transfer modeling along with experimental tests. Analytical solutions for the film thickness and the lengthwise vapor temperature distribution for a single heat pipe are derived. Experimental tests on single radially rotating miniature heat pipes and sector heat pipes are undertaken with different important parameters and the manner in which these parameters affect heat pipe operation. Analytical and experimental studies have proven that the radially rotating miniature heat pipes have an incredibly high effective thermal conductance and an enormous heat transfer capability. Concurrently, the heat pipe has an uncomplicated structure and relatively low manufacturing costs. The heat pipe can also resist strong vibrations and is well suited for a high temperature environment. Hence, the heat pipes with a common reservoir make incorporation of heat pipes into turbo-machinery much more feasible and cost effective.

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Nickol, Jeremy B. "Heat Transfer Measurements and Comparisons for a Film Cooled Flat Plate with Realistic Hole Pattern in a Medium Duration Blowdown Facility." The Ohio State University, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=osu1365421507.

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Agricola, Lucas. "Nozzle Guide Vane Sweeping Jet Impingement Cooling." The Ohio State University, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=osu1525436077557298.

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Nickol, Jeremy B. "Airfoil, Platform, and Cooling Passage Measurements on a Rotating Transonic High-Pressure Turbine." The Ohio State University, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=osu1459857581.

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Janakiraman, S. V. "Fluid flow and heat transfer in transonic turbine cascades." Thesis, This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-06112009-063614/.

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Scheepers, Gerard. "An experimental and numerical study of heat transfer augmentation near the entrance to a film cooling hole." Pretoria : [s.n.], 2007. http://upetd.up.ac.za/thesis/available/etd-08272008-163851/.

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Books on the topic "Heat-transfer; Aerospace gas turbines"

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Garg, Vijay K. Heat transfer in gas turbines. [Cleveland, Ohio]: National Aeronautics and Space Administration, Glenn Research Center, 2001.

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2

North Atlantic Treaty Organization. Advisory Group for Aerospace Research and Development. Heat Transfer and Cooling in Gas Turbines. S.l: s.n, 1985.

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Meeting, American Society of Mechanical Engineers Winter. Heat transfer in gas turbine engines. New York, N.Y. (345 E. 47th St., New York 10017): The Society, 1987.

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4

1953-, Dutta Sandip, and Ekkad Srinath 1958-, eds. Gas turbine heat transfer and cooling technology. New York: Taylor & Francis, 2000.

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International Symposium on Heat Transfer in Turbomachinery (1992 Marathon, Greece). Heat transfer in turbomachinery: Proceedings of the International Symposium on Heat Transfer the (sic) in Turbomachinery, Greece, August 1992. New York: Begell House, 1994.

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Sandīpa, Datta, and Ekkad Srinath 1958-, eds. Gas turbine heat transfer and cooling technology. 2nd ed. Boca Raton, FL: Taylor & Francis, 2012.

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7

American Society of Mechanical Engineers. Heat transfer in gas turbine engines and three-dimensional flows. New York: American Society of Mechanical Engineers, 1988.

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American Society of Mechanical Engineers. Winter Meeting. Heat transfer in gas turbine engines - 1991: Presented at the Winter Annual Meeting of the American Society of Mechanical Engineers, Atlanta, Georgia, December 1-6, 1991. New York, N.Y: ASME, 1991.

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9

Ameri, A. A. Analysis of gas turbine rotor blade tip and shroud heat transfer. [Washington, DC: National Aeronautics and Space Administration, 1996.

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Ameri, A. A. Analysis of gas turbine rotor blade tip and shroud heat transfer. [Washington, DC: National Aeronautics and Space Administration, 1998.

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Book chapters on the topic "Heat-transfer; Aerospace gas turbines"

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Sreenath, P., Shambhoo, H. S. Raghukumar, C. Rajashekar, A. Davis, and J. J. Isaac. "Triggering of Flow Instabilities by Simulated Sub/Supercritical Rayleigh Heat Addition in an Aero-Gas Turbine Afterburner." In Proceedings of the National Aerospace Propulsion Conference, 203–15. Singapore: Springer Singapore, 2020. http://dx.doi.org/10.1007/978-981-15-5039-3_12.

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Sunden, B., J. Larocque, and Z. Wu. "Numerical Simulation of Heat Transfer from Impinging Swirling Jets." In Impingement Jet Cooling in Gas Turbines, 185–202. WIT Press, 2014. http://dx.doi.org/10.2495/978-1-84564-906-7/007.

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"Convective heat transfer in blade cooling and heat-exchanger design." In The Design of High-Efficiency Turbomachinery and Gas Turbines. The MIT Press, 2014. http://dx.doi.org/10.7551/mitpress/9940.003.0017.

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Takeishi, K., and Y. Oda. "Experimental and Numerical Study on Heat Transfer Enhancement of Impingement Jet Cooling by Adding Ribs on Target Surface." In Impingement Jet Cooling in Gas Turbines, 203–31. WIT Press, 2014. http://dx.doi.org/10.2495/978-1-84564-906-7/008.

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Conference papers on the topic "Heat-transfer; Aerospace gas turbines"

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Thakur, S., and J. Wright. "Conjugate heat transfer in a gas turbine blade trailing-edge cavity." In 40th AIAA Aerospace Sciences Meeting & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2002. http://dx.doi.org/10.2514/6.2002-496.

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Zhou, J. M., M. Robichaud, W. Habashi, and W. Ghaly. "CFD predictions of flow and heat transfer in the coolant passages of gas turbine blades." In 34th Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1996. http://dx.doi.org/10.2514/6.1996-450.

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Amano, Ryoichi S., and Saman Beyhaghi. "Heat Transfer in a Rotating Two-pass Square Channel Representing Internal Cooling of Gas Turbine Blades." In 54th AIAA Aerospace Sciences Meeting. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2016. http://dx.doi.org/10.2514/6.2016-0655.

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Stuttaford, Peter, Philip Rubini, Peter Stuttaford, and Philip Rubini. "Assessment of a radiative heat transfer model for a new gas turbine combustor preliminary design tool." In 35th Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1997. http://dx.doi.org/10.2514/6.1997-294.

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Barger, Ben, Chao Moua, Ryan Holschbach, Justin Wood, Lee Wisinski, Sourabh Kumar, and Ryoichi Amano. "Experimental Analysis on Heat Transfer Distributions Corresponding to Varying Internal Geometries in Gas Turbine Blade Cooling Systems." In 50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2012. http://dx.doi.org/10.2514/6.2012-1003.

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Giridhara Babu, Yepuri, Gururaj Lalgi, Ashok Babu Talanki Puttarangasetty, Jesuraj Felix, Sreenivas Rao V. Kenkere, and Nanjundaiah Vinod Kumar. "Experimental and Numerical Investigation of Effect of Blowing Ratio on Film Cooling Effectiveness and Heat Transfer Coefficient Over a Gas Turbine Blade Leading Edge Film Cooling Configurations." In ASME 2013 Gas Turbine India Conference. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/gtindia2013-3552.

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Film cooling is one of the cooling techniques to cool the hot section components of a gas turbine engines. The gas turbine blade leading edges are the vital parts in the turbines as they are directly hit by the hot gases, hence the optimized cooling of gas turbine blade surfaces is essential. This study aims at investigating the film cooling effectiveness and heat transfer coefficient experimentally and numerically for the three different gas turbine blade leading edge models each having the one row of film cooling holes at 15, 30 and 45 degrees hole orientation angle respectively from stagnation line. Each row has the five holes with the hole diameter of 3mm, pitch of 20mm and has the hole inclination angle of 20deg. in spanwise direction. Experiments are carried out using the subsonic cascade tunnel facility of National Aerospace Laboratories, Bangalore at a nominal flow Reynolds number of 1,00,000 based on the leading edge diameter, varying the blowing ratios of 1.2, 1.50, 1.75 and 2.0. In addition, an attempt has been made for the film cooling effectiveness using CFD simulation, using k-€ realizable turbulence model to solve the flow field. Among the considered 15, 30 and 45 deg. models, both the cooling effectiveness and heat transfer coefficient shown the increase with the increase in hole orientation angle from stagnation line. The film cooling effectiveness increases with the increase in blowing ratio upto 1.5 for the 15 and 30 deg. models, whereas on the 45 deg. model the increase in effectiveness shown upto the blowing ratio of 1.75. The heat transfer coefficient values showed the increase with the increase in blowing ratio for all the considered three models. The CFD results in the form of temperature, velocity contours and film cooling effectiveness values have shown the meaningful results with the experimental values.

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Constantin, Sandu, and Dan Brasoveanu. "Relative Cooling With Liquid for Gas Turbines: Part I — A Solution for Exceeding 2000 K at Turbine Inlet." In ASME 2001 International Design Engineering Technical Conferences and Computers and Information in Engineering Conference. American Society of Mechanical Engineers, 2001. http://dx.doi.org/10.1115/detc2001/cie-21763.

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Abstract The thermal efficiency of gas turbines is critically dependent on the temperature of burnt gases at turbine inlet, the higher this temperature the higher the efficiency. Stochiometric combustion would provide maximum efficiency, but in the absence of an internal cooling system, turbine blades cannot tolerate gas temperatures that exceed 1300 K. Therefore, for this temperature, the thermal efficiency of turbine engine is 40% less than theoretical maximum. Conventional air-cooling techniques of turbine blades allow inlet temperatures of about 1500 K on current operating engines yielding thermal efficiency gains of about 6%. New designs, that incorporate advanced air-cooling methods allows inlet temperatures of 1750–1800 K, with a thermal efficiency gain of about 6% relative to current operating engines. This temperature is near the limit allowed by air-cooling systems. Turbine blades can be cooled with air taken from the compressor or with liquid. Cooling systems with air are easier to design but have a relatively low heat transfer capacity and reduce the efficiency of the engine. Some cooling systems with liquid rely on thermal gradients to promote re-circulation from the tip to the root of turbine blades. In this case, the flow and cooling of liquid are restricted. For best results, cooling systems with liquid should use a pump to re-circulate the coolant. In the past, designers tried to place this pump on the engine stator and therefore were unable to avoid high coolant losses because it is impossible to reliably seal the stator-rotor interface. Therefore it was assumed that cooling systems with liquid could not incorporate pumps. This is an unwarranted assumption as shown studying the system in a moving frame of reference that is linked to the rotor. Here is the crucial fact overlooked by previous designers. The relative motion of engine stator with respect to the rotor is sufficient to motivate a cooling pump. Both the pump and heat exchange system that is required to provide rapid cooling of liquid with cold ambient air, could be located within the rotor. Therefore, the entire cooling system can be encapsulated within the rotor and the sealing problem is circumvented. Compared to recent designs that use advanced air-cooling methods, such a liquid cooling system would increase the thermal efficiency by 8%–11% because the temperatures at turbine inlet can reach stoichiometric levels and most of the heat extracted from turbine during cooling is recuperated. The appreciated high reliability of the system will permit a large applicability in aerospace propulsion.

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Krishnanunni, Abhay Girija, Jayakumar Janardhanan Sarasamma, Yepuri Giridhara Babu, and Felix Jesuraj. "CFD Simulation of Open-Cell Aluminum Metal Foams for Pressure Drop Characterization." In ASME 2013 Gas Turbine India Conference. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/gtindia2013-3605.

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Many researchers have done the CFD simulation of open-cell aluminum metal foams with a unit cell with periodic boundary conditions. However, this does not represent a real life situation, as the foam-fluid interactions cannot be properly modeled. In the present study the simulation is done for metal foam with the more number of foam cells to proximate the conditions close to the actual situations. The CFD simulation of open-cell aluminum metal foams was done using ANSYS FLUENT. The results are obtained by solving the Continuity, Momentum and Energy equations and standard k-ε turbulence model is used for simulation. The boundary conditions applied are same as those applied during the experiments conducted at Heat Transfer Lab, National Aerospace Laboratories, Bangalore. In this study the Aluminum Alloy (Al 6101-T6) metal foam of pore density 10 ppi is used for CFD analysis.

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McDonald, Colin F. "Helium and Combustion Gas Turbine Power Conversion Systems Comparison." In ASME 1995 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1995. http://dx.doi.org/10.1115/95-gt-263.

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For closed-cycle gas turbines, in a size to meet utility power generation needs, the selection of helium as the working fluid represents the best solution in terms of the overall power conversion system considering the differing requirements of the turbomachinery and heat exchangers. Helium is well suited for the nuclear Brayton cycle because it is neutronically inert. The impact of helium’s unique properties on the performance and size of the power conversion system components is discussed in this paper. The helium gas turbine plants, that have operated were based on 1950s and 1960s technology, represent a valuable technology base in terms of practical experience gained. However, the design of the Gas Turbine Modular Helium Reactor (GT-MHR), which could see utility service in the first decade of the 21st century will utilize turbomachinery and heat exchanger technologies from the combustion gas turbine and aerospace industries. An understanding of how the design of power conversion systems for closed-cycle plants and combustion gas turbines are affected by the working fluids (i.e., helium and air, respectively) is the major theme of this paper.

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Herring, Neal R., and Stephen D. Heister. "Review of the Development of Compact, High Performance Heat Exchangers for Gas Turbine Applications." In ASME 2006 International Mechanical Engineering Congress and Exposition. ASMEDC, 2006. http://dx.doi.org/10.1115/imece2006-14920.

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Abstract:

This study provides a review of the current state-of-the-art in compact heat exchangers and their application to gas turbine thermal management. Specifically, the challenges and potential solutions for a cooled cooling air system using the aircraft fuel as a heat sink were analyzed. As the sensible heat absorbed by the fuel in future engines is increased, the fuel will be exposed to increasingly hotter temperatures. This poses a number of design challenges for fuel-air heat exchangers. The most well known challenge is fuel deposition or coking. Another problem encountered at high fuel temperatures is thermo-acoustic oscillations. Thermo-acoustic oscillations have been shown to occur in many fluids when heated near the critical point, yet the mechanism of these oscillations is poorly understood. In some cases these instabilities have been strong enough cause failure in the thin walled tubes used in heat exchangers. For the specific application of a fuel-air heat exchanger, the advantages of a laminar flow device are discussed. These devices make use of the thermal entry region to achieve high heat transfer coefficients. To increase performance further, heat transfer enhancement techniques were reviewed and the feasibility for aerospace heat exchangers was analyzed. Two of the most basic techniques for laminar flow enhancement include tube inserts and swirl flow devices. Additionally, the effects of these devices on both coking and instabilities have been assessed.

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