Relevant bibliographies by topics / Gas turbine swirl injector

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Academic literature on the topic 'Gas turbine swirl injector'

Author: Grafiati
Published: 6 September 2023

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Journal articles on the topic "Gas turbine swirl injector"

1

Pham, Vu Thanh Nam. "AN IMAGE PROCESSING APPROACH FOR DETERMINING THE SPRAY CONE ANGLE OF A PRESSURE SWIRL INJECTOR EQUIPPED IN A GAS-TURBINE ENGINE." Journal of Science and Technique 16, no. 2 (2022): 33–47. http://dx.doi.org/10.56651/lqdtu.jst.v16.n02.265.

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This paper adopts the directionality tool provided by the ImageJ package to determine the spray cone angle of a gas-turbine engine’s injector. An imaging experiment system has been developed in this study to image a spray of a practical gas-turbine injector under injection pressure conditions varying from 2 to 6 bars. The results show that the reliability of the measurement is achieved when analyzing at least 500 images. Preferably, using 1500 images shows the uncertainty of less than 0.5% (approximately corresponding with 0.2° of the angle). The average spray cone angle varies between 100° and 128.15° when the injection pressure increases from 2 to 6 bars. An accurate determination of the spray cone angle helps to improve the quality of research on micro and macro spray characteristics including the droplet concentration and distribution. The results could also be utilized to develop a diagnostic technique for gas-turbine injectors.
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2

McGuirk, J. J. "The aerodynamic challenges of aeroengine gas-turbine combustion systems." Aeronautical Journal 118, no. 1204 (2014): 557–99. http://dx.doi.org/10.1017/s0001924000009386.

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Abstract The components of an aeroengine gas-turbine combustor have to perform multiple tasks – control of external and internal air distribution, fuel injector feed, fuel/air atomisation, evaporation, and mixing, flame stabilisation, wall cooling, etc. The ‘rich-burn’ concept has achieved great success in optimising combustion efficiency, combustor life, and operational stability over the whole engine cycle. This paper first illustrates the crucial role of aerodynamic processes in achieving these performance goals. Next, the extra aerodynamic challenges of the ‘lean-burn’ injectors required to meet the ever more stringent NO x emissions regulations are introduced, demonstrating that a new multi-disciplinary and ‘whole system’ approach is required. For example, high swirl causes complex unsteady injector aerodynamics; the threat of thermo-acoustic instabilities means both aerodynamic and aeroacoustic characteristics of injectors and other air admission features must be considered; and high injector mass flow means potentially strong compressor/combustor and combustor/turbine coupling. The paper illustrates how research at Loughborough University, based on complementary use of advanced experimental and computational methods, and applied to both isolated sub-components and fully annular combustion systems, has improved understanding and identified novel ideas for combustion system design.
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3

So, Younseok, Yeoungmin Han, and Sejin Kwon. "Combustion Characteristics of Multi-Element Swirl Coaxial Jet Injectors under Varying Momentum Ratios." Energies 14, no. 13 (2021): 4064. http://dx.doi.org/10.3390/en14134064.

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The combustion characteristics of a staged combustion cycle engine with an oxidizer-rich preburner were experimentally studied at different momentum ratios of multi-element injectors. Propellants were simultaneously supplied as a liquid–liquid–liquid system, and an injector was designed in which a swirl coaxial jet is sprayed. The injector burned the propellants in the inner chamber which had a temperature greater than 2000 K. To cool the combustion gas, a liquid oxidizer was supplied to the cooling channel outside the injector. To prevent the turbine blades from melting, the temperature of the combustion gas was maintained below 700 K. To confirm the combustion characteristics at different momentum ratios of the high-temperature combustion gas inside the injector and the low-temperature liquid oxidizer outside the injector, three types of injectors were designed and manufactured with different momentum ratios: MR 3.0, MR 3.3, and MR 3.7. In this study, the results of the combustion test for each type were compared for 30 s. For ORPB-A, a combustion pressure of 18.5 MPaA, fuel mass flow rate of 0.26 kg/s, oxidizer mass flow rate of 15.3 kg/s, and turbine inlet temperature of 686 K were obtained in the combustion stability period of 29.0–29.5 s. The combustion efficiency was 98% for MR 3.0 (ORPB-A), which was superior to that for other momentum ratios. In addition, during the combustion test for MR 3.0, the fluctuations in the characteristic velocity, combustion pressure, and propellant mass flow rate were low, indicating that combustion was stable. The three types of combustion instability were all less than 0.8%, thus confirming that the combustion stability was excellent.
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4

WANG, SHANWU, VIGOR YANG, GEORGE HSIAO, SHIH-YANG HSIEH, and HUKAM C. MONGIA. "Large-eddy simulations of gas-turbine swirl injector flow dynamics." Journal of Fluid Mechanics 583 (July 4, 2007): 99–122. http://dx.doi.org/10.1017/s0022112007006155.

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A comprehensive study on confined swirling flows in an operational gas-turbine injector was performed by means of large-eddy simulations. The formulation was based on the Favre-filtered conservation equations and a modified Smagorinsky treatment of subgrid-scale motions. The model was then numerically solved by means of a preconditioned density-based finite-volume approach. Calculated mean velocities and turbulence properties show good agreement with experimental data obtained from the laser-Doppler velocimetry measurements. Various aspects of the swirling flow development (such as the central recirculating flow, precessing vortex core and Kelvin–Helmholtz instability) were explored in detail. Both co- and counter-rotating configurations were considered, and the effects of swirl direction on flow characteristics were examined. The flow evolution inside the injector is dictated mainly by the air delivered through the primary swirler. The impact of the secondary swirler appears to be limited.
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5

Woo, Seongphil, Jungho Lee, Yeoungmin Han, and Youngbin Yoon. "Experimental Study of the Combustion Efficiency in Multi-Element Gas-Centered Swirl Coaxial Injectors." Energies 13, no. 22 (2020): 6055. http://dx.doi.org/10.3390/en13226055.

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The effects of the momentum-flux ratio of propellant upon the combustion efficiency of a gas-centered-swirl-coaxial (GCSC) injector used in the combustion chamber of a full-scale 9-tonf staged-combustion-cycle engine were studied experimentally. In the combustion experiment, liquid oxygen was used as an oxidizer, and kerosene was used as fuel. The liquid oxygen and kerosene burned in the preburner drive the turbine of the turbopump under the oxidizer-rich hot-gas condition before flowing into the GCSC injector of the combustion chamber. The oxidizer-rich hot gas is mixed with liquid kerosene passed through combustion chamber’s cooling channel at the injector outlet. This mixture has a dimensionless momentum-flux ratio that depends upon the dispensing speed of the two fluids. Combustion tests were performed under varying mixture ratios and combustion pressures for different injector shapes and numbers of injectors, and the characteristic velocities and performance efficiencies of the combustion were compared. It was found that, for 61 gas-centered swirl-coaxial injectors, as the moment flux ratio increased from 9 to 23, the combustion-characteristic velocity increased linearly and the performance efficiency increased from 0.904 to 0.938. In addition, excellent combustion efficiency was observed when the combustion chamber had a large number of injectors at the same momentum-flux ratio.
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6

Lezsovits, Ferenc, Sándor Könczöl, and Krisztián Sztankó. "CO emission reduction of a HRSG duct burner." Thermal Science 14, no. 3 (2010): 845–54. http://dx.doi.org/10.2298/tsci1003845l.

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A heat-recovery steam generator was erected after a gas-turbine with a duct burner into the district heat centre. After commissioning, the CO emissions were found to be above the acceptable level specified in the initial contract. The Department of Energy Engineering of the BME was asked for their expert contribution in solving the problem of reducing these CO emissions. This team investigated the factors that cause incomplete combustion: the flue-gas outlet of the gas-turbine has significant swirl and rotation, the diffuser in between the gas-turbine and heat-recovery steam generator is too short and has a large cone angle, the velocity of flue-gas entering the duct burner is greater than expected, and the outlet direction of the flammable mixture from the injector of the duct burner was not optimal. After reducing the flow swirl of flue-gas and modifying the nozzle of the duct burner as suggested by the Department of Energy Engineering of the BME, CO emissions have been reduced to an acceptable level. The method involves the application of CFD modeling and studying images of the flames which proved to be very informative.
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7

Durbin, M. D., M. D. Vangsness, D. R. Ballal, and V. R. Katta. "Study of Flame Stability in a Step Swirl Combustor." Journal of Engineering for Gas Turbines and Power 118, no. 2 (1996): 308–15. http://dx.doi.org/10.1115/1.2816592.

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A prime requirement in the design of a modern gas turbine combustor is good combustion stability, especially near lean blowout (LBO), to ensure an adequate stability margin. For an aeroengine, combustor blow-off limits are encountered during low engine speeds at high altitudes over a range of flight Mach numbers. For an industrial combustor, requirements of ultralow NOx emissions coupled with high combustion efficiency demand operation at or close to LBO. In this investigation, a step swirl combustor (SSC) was designed to reproduce the swirling flow pattern present in the vicinity of the fuel injector located in the primary zone of a gas turbine combustor. Different flame shapes, structure, and location were observed and detailed experimental measurements and numerical computations were performed. It was found that certain combinations of outer and inner swirling air flows produce multiple attached flames, aflame with a single attached structure just above the fuel injection tube, and finally for higher inner swirl velocity, the flame lifts from the fuel tube and is stabilized by the inner recirculation zone. The observed difference in LBO between co- and counterswirl configurations is primarily a function of how the flame stabilizes, i.e., attached versus lifted. A turbulent combustion model correctly predicts the attached flame location(s), development of inner recirculation zone, a dimple-shaped flame structure, the flame lift-off height, and radial profiles of mean temperature, axial velocity, and tangential velocity at different axial locations. Finally, the significance and applications of anchored and lifted flames to combustor stability and LBO in practical gas turbine combustors are discussed.
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8

Sung, Hong-Gye. "Combustion dynamics in a model lean-premixed gas turbine with a swirl stabilized injector." Journal of Mechanical Science and Technology 21, no. 3 (2007): 495–504. http://dx.doi.org/10.1007/bf02916311.

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9

Mardani, Amir, Rezapour Rastaaghi, and Fazlollahi Ghomshi. "Liquid petroleum gas flame in a double-swirl gas turbine model combustor: Lean blow-out, pollutant, preheating." Thermal Science, no. 00 (2020): 139. http://dx.doi.org/10.2298/tsci190623139m.

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In this paper, lean blow-out (LBO) limits in a double swirl gas turbine model combustor were investigated experimentally for Liquid Petroleum Gas (LPG) fuel. The LBO curve was extracted for different combustor configurations. While burner could operate reasonably under ultra-lean conditions, two different sets of operating conditions, one with a low flow rate (LFR) and another one with high flow rate (HFR), are identified and studied in terms of LBO and pollutant. Results showed that while the flame structure was similar in both cases, the chamber responses to geometrical changes and also preheating are minimal at the LFR. That means confinement and injector type have desirable effects on stability borders but not for the LFR. The channeled injector shifted down the LBO limit around 28 percent at HFR. Measurements on the combustor exhaust gas composition and temperature indicate a region with relatively complete combustion and reasonable temperature and a very low level of exhaust NOx pollutants (i.e., below ten ppm) at about 25-50% above the LBO. In this operating envelope, a burner power increment led to a higher exhaust average temperature and combustion efficiency, while NOx formation decreased. Preheating the inlet air up to 100?C results in an improvement in burner stability in about 10 percent, but NOx production intensifies more than three times. Results indicate that the LBO limit is configured more by the burner design and aerodynamic aspects rather than the fuel type.
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10

Cheng, R. K., D. Littlejohn, P. A. Strakey, and T. Sidwell. "Laboratory investigations of a low-swirl injector with H2 and CH4 at gas turbine conditions." Proceedings of the Combustion Institute 32, no. 2 (2009): 3001–9. http://dx.doi.org/10.1016/j.proci.2008.06.141.

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