• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.353-358
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• 2009

For the successful development of 75 tonf class liquid rocket engine, a plenty of tests on each engine component has to be performed and this is equally true for a combustor. However the test facility which is in operation at Korea Aerospace Research Institute lacks its capacity to perform fire tests of a 75 tonf class combustor at its nominal thrust. Since the test facility has to be ready prior to the start of development tests, it is very urgent to establish the test facility. The preliminary design of a test facility for a 75 tonf class combustor which was performed according to the urgent necessity is described in the paper.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.506-510
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• 2008

Ramjets are capable of much higher specific impulse than liquid rocket engines for high speed flight in the atmosphere. Ramjets, however, cannot generate thrust at low flight speed. Therefore, an additional propulsion device to accelerate the ramjet vehicle to a supersonic speed is required. In this study, we propose a novel ramjet propulsion system with a $H_2O_2$/Kerosene rocket as the accelerator for initial stage. In order to test the feasibility of this concept, consecutive reactors was built; one for the decomposition of $H_2O_2$ and the other for kerosene combustion. Decomposed $H_2O_2$ jet was injected to combustor through converging nozzle from gas generator and over this hot oxygen jet, kerosene was injected by spay injector. Through the various test cases, hypergolic ignition test was carried out and steady combustion was achieved.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.461-464
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• 2012

The high-frequency data acquisition system of the rocket engine test facility has been updated to perform hot-firing tests of 7 ton-class liquid rocket engine combustion chambers which will be used for the third stage of the Korea space launch vehicle II. The paper deals with the design of the updated high-frequency data acquisition system and explains its main functions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.180-185
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• 2004

Our previous study showed that although the hybrid rocket engine with swirling gaseous oxygen had high performance, a direct injection of LOX with swirl into the combustion chamber of the hybrid rocket engine lowered the performance of the engine, compared to that with gaseous oxygen. In order to clarify this reason, combustion tests of a small PMMA combustor with an inner port diameter of 2 mm were conducted in liquid oxygen flow by comparison with gaseous oxygen flow. Although the oxygen mass fluxes of LOX were about two orders of magnitude larger than those of gaseous oxygen, the fuel regression rate of LOX were remarkably smaller than those of gaseous oxygen. For both liquid and gaseous oxygen, diffusion flames in the port of the grain controlled the combustion process of PMMA in oxygen flow. These results may be explained by the fact that only small amount of LOX vaporized and consumed in the combustion with PMMA while flowing through the port due to relatively larger latent heat of injected liquid oxygen compared to the heat of release by combustion which depended on the burning surface area of PMMA.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.191-195
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• 2006

It is definitely required to control dispersion of the rocket engine performance in order to accomplish the mission of a launch vehicle successfully. A performance dispersion analysis was conducted for a gas generator cycle liquid rocket engine and the required pressure drops were estimated for engine tunning. As a result, the vacuum thrust dispersion of the engine was from +9.1% to -8.7% and the mixture ratio deviated from +9.7% to -9.6% from the nominal value due to the errors of components and the engine inlet condition of propellants. The required pressure drop in the LOx line to the combustor is higher than in the fuel line for same mixture ratio change.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.8
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• pp.704-710
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• 2017

The supply of the liquid oxygen into a rocket combustor is simply controlled by the 'on' and 'off' positions of a main oxidizer shut-off valve. However, the partially opened position of a three-position valve can control and optimize the engine start transients by regulating the liquid oxygen flow rate during the start-up of the engine. In this paper, the design and performances of a three-position pneumatic poppet valve, which is intended to be employed in liquid rocket engines, have been presented.

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.195-204
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• 2006

A linear acoustic analysis has been performed to elucidate damping characteristics of acoustic cavities in a liquid rocket combustor. Results have shown that resonant frequencies of acoustic cavity obtained by classical theoretic approach and by the present linear analysis are somewhat different with each other. This difference is attributed to the limitation of the simplified classical theory. To quantify the damping characteristics, acoustic impedance has been introduced and resultant absorption coefficient and conductance have been evaluated. Satisfactory agreement has been achieved with previous experiment. Finally the design procedure for an optimal tuning of acoustic cavity has been established.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.259-264
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• 2004

The critical component of combustor having high combustion efficiency for high performance liquid rocket engine is injector. The results of design and hot firing tests of six sub-scale combustors which have respectively an impinging type injector(1ea.), an bi-propellant swirl closed injector(1ea.), and hi-propellant swirl mixed injector(4ea.) were described in this paper. The combustion test were successfully performed. The combustion efficiency have higher value than predicted value and high frequency combustion instability does not occur.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.305-308
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• 2005

This paper presents a fabrication of a full-stale combustion chamber of a liquid rocket engine for a ground hot firing test. Engine drawings for manufacturing were prepared after conceptual and detail designs. The combustor is composed of a head and a chamber. SUS316L is used for materials of the head because of the good quality in low temperature. Inner materials of the ablative cooling chamber is silica/phenolic and outer case materials is the SUS316L. Materials of the regenerative cooling chamber are C18200 and SUS316L. After lathe, general milling and MCT machinings, components were finished by electrolytic polishing. A brazing method was applied for bonding the injectors and the injector plate, the regenerative cooling chamber because of structure configurations.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.8
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• pp.69-76
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• 2003

For stable combustion and safety of a structure of the propulsion system, a cooling system to the liquid rocket engine should be incorporated. In this study, Eulerian-Lagrangian scheme for two phase combustion, nongray radiation and soot formation effect, and film-wall interaction have been introduced to study the effect of film cooling. After briefly introducing the governing equation, combustion characteristics with change of wall temperature has been investigated by varying such parameters as fuel mass fraction for film cooling, diameter of the fuel droplet, overall mixture fraction of oxygen to fuel. Also, radiative heat flux is compared with the conductive one at the combustor wall.