• Journal of ILASS-Korea
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• v.16 no.4
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• pp.195-200
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• 2011

The pressure-swirl atomizer is widely used for the injectors in liquid rocket engines thanks to its high performance atomization and broad stability margin range. Spray mixed-interaction is an important area of study especially in cases where the propellant is mixed by spray interaction after an oxidant and a fuel are discharged separately. This interaction of sprays results in a significant modification of the spray characteristics such as the spatial evolution of the sprays. Experiments are conducted by a photographic technique to quantify the merged angle of the interaction regions of sprays from two pressure-swirl injectors. The experimental results show that the merged angle is mainly determined by the momentum flux ratios between two swirled sprays.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.24 no.11
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• pp.1457-1463
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• 2000

Experimental studies on determination of the supply leading time of propellants to combustion chamber have been made to stably and efficiently guarantee the ignitions process with liquid rocket engine. The propellant used is a Kerosene as fuel and a liquid oxygen as oxidizer. FOOF type of three injectors are set with an angle of 135。 and the combustion chamber pressure is 200psi. The present experiment program also includes the stability on the quadlet type of ignitor using the triehylaluminum (TEAL) as an ignition source. Experimental results clarifies that the propellant supply through LOx leading to combustion chamber is proper for stable ignition and combustion processes based on the fuel and oxidizer manifold pressures, combustion chamber pressure, and the variation of flame length from the nozzle exit with lapse time, and shows that the leading supply time pf propellants effects the engine performance little.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.4
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• pp.48-54
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the lab-scale liquid rocket engine using liquid oxygen(LOx) and Jet A-1(Jet engine fuel) as propellants. Film coolants(Jet A-1 and water) was injected through the film cooling injector. The outside wall temperature of the combustor and film cooled length were determined for chamber pressure, mixture ratio, and the different geometries(injection angle) with the percent film coolant flow rate. The loss of characteristic velocity was determined for the case of film cooling with water and Jet A-1. As chamber pressure increased, the outside wall temperature increased in the nozzle but unchanged over the 9 percent film coolant flow rate for the combustion chamber used in this study. Characteristic velocity wasn't affected with the mixture ratio over the 9 percent film coolant flow rate.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.11
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• pp.1140-1147
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• 2009

Flow analyses have been performed to investigate the uniformity of propellant flow through the fuel and oxidizer manifolds of a full-scaled gas generator for a pump-fed liquid rocket engines. Injectors were simulated as porous medium layers having equivalent pressure drops. The uniformity of propellants has been analyzed for 3 fuel rings and 3 injector head configurations. The mixture ratio distribution at the exit of injectors has been estimated from the mass flow rates of fuel and oxidizer. The best configuration of fuel ring and injection head was selected through these flow analyses.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.353-357
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• 2005

LOx post damage occurs from the development process of the full-scale gas generator which is necessary to 30 tonf class engine development was described. The cause and analysis for damage was described. The combustion test result of 4 injector, the full-scale gas generator and redesigned injector was described. Combustion instability, purge, the low momentum of LOx spray, small recess number, the low flow of LOx, and the high spray angle is main reason the possibility of knowing. The redesign for the injector in the direction of increase of recess number, increase of LOx and fuel spray angle, decrease of gap interval between the LOx post outer wall and fuel screen and increase of LOx post wall thick became accomplished.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.10
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• pp.1038-1047
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• 2009

The development process of a gas generator for a 30-tonf pump-fed space liquid rocket engine is described. Starting from the development of an injector, followed by subscale and full-scale test specimens, the development of LOx/kerosene fuel-rich gas generator has been concluded successfully. Various analytical methods have been utilized in the course of design and the performance requirements have been verified experimentally through ignition tests, combustion performance and stability assessment tests and duration tests. The gas generator has proven its workability and stability within a defined operation window of varying chamber pressure and mixture ratio and demonstrated compliance to the performance and life time requirements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.69-72
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• 2011

Hot-firing tests were performed on the third sub-scale gas generator for development of a 75 ton-class liquid rocket engine. This paper deals with the analysis results of low-frequency combustion instabilities that were encountered during combustion tests of the gas generator. The low-frequency pressure fluctuations seem to be related to chamber pressure and pressure drops through oxidizer/fuel injectors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.175-178
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• 2006

A numerical study for spray combustion of fluctuated fuel and oxidizer droplets injected into combustion chamber has been conducted for the analysis of spray combustion considering characteristics of injector. The 2 dimensional unsteady state flow fields have been calculated by using QUICK Scheme and SIMPLER Algorithm. As the spray model, DSF model and Euler-Lagrange Scheme have been used. The sine Auction has been used for droplet fluctuation model of fuel and oxidizer, while the coupling effects of the droplets between gas phase and evaporated vapor have been calculated by using PSIC model.

• Proceedings of the KSME Conference
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• 2002.08a
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• pp.255-258
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• 2002

Liquid rocket engines using liquefied natural gas (LNG) or methane as a fuel is known to have several good characteristics, such as high specific impulse compared to other hydrocarbon fuels, environment-friendly exhaust gas, low production cost, and re-usability with low soot generation in the cooling channel. In this study, experimental combustion chambers capable of using LNC and $CH_{4}$ are being researched through experimental firing tests, and within easy range of eyes' inspection, there are the periodical existence of soot or discoloration in the chamber wall surface. This result means that mixture ratio of oxidizer and fuel fluctuates periodically between outer-row injectors in the mixing head in the circumferential direction. Therefore, based on this phenomenon, the variation of mixture ratio near the chamber wall caused by the spill pattern of a shear coaxial injector was analyzed quantitatively and the thermal heat flux Into the cooling channel is modified. Then, the calculated and modified results are compared with the measured ones.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.346-350
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• 2005

Effects of the changes of a film cooling mass flow rate and operating conditions on the heat flux characteristics of the subscale calorimeter were studied. A film cooling ring with twelve orifices is inserted between the injector head and the calorimeter. The calorimeter is composed of nineteen cooling channels. When a mass flow rate of film cooling is 10.5 % of a main fuel mass flow rate, maximum heat flux at the nozzle throat is decreased by 30% compared to that without film cooling. In the OD3(of-design point) test result, maximum heat flux at the nozzle throat is increased by 31% compared to that of the DP(design point) test when a film cooling flow rate is zero.