• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.370-373
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• 2007

In this study, a hybrid rocket motor with separable and detachable oxidizer tank from combustion chamber is developed. Initially, the measured thrust of the motor showed about 30% of the design thrust since the oxidizer supply was not enough. In order to solve this problem, application is made to expand the orifice diameter of oxidizer injector empirically, so that the mass flow rate of oxidizer was improved. The improved performance was about 60% of design thrust, 18kgf, and thrust-to-weight ratio was reasonable, compared with other sounding rockets.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.4
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• pp.91-101
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• 2004

With the mathematic linear model of a combustor which consists of a combustion chamber and injectors, the analysis of low frequency dynamic characteristics of a liquld-propellant rocket engine combustor was performed. Propellant mass flowrate was varied by combustion chamber pressure feedback, therefore low frequency oscillation was appeared. Increasing the time constant of a combustion chamber and injector pressure differences and decreasing combustion time delay increased the combustor system stability. The variation of injector time constant little affected stability. The system was always stable, when there was no combustion time delay. Increasing combustion time delay decreased oscillation frequency and damping ratio, and the system eventually became unstable.

• Proceedings of the KSME Conference
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• 2002.08a
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• pp.283-286
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• 2002

The mixing of fuel and oxidizer and each mass distribution of unlike split triplet(F-O-O-F) injector for liquid rocket which are known to affect the combustion efficiency significantly, has been investigated using PLLIF technique. Experiment is conducted to investigate the effect of mixture ratio(MiR), momentum ratio(MoR) and impinging angle on mixing efficiency. The mixing efficiency, which is introduced by Rupe, gives the global chracteristic of spray mixing. Experiment far comparison with triplet injector(F-O-F), which has same momentum ratio and impinging angle with split triplet injector is conducted.

• Journal of ILASS-Korea
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• v.7 no.1
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• pp.21-28
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• 2002

The mixing of propellant and its mass distribution of unlike-doublet impinging injector, which is known to affect the combustion efficiency significantly, have been studied using PLIF(Planar Laser Induced Fluorescence). The results show that fuel jet penetrates considerably into the oxidizer jet at impinging point as variation of momentum ratio. and then stream flows inclined because of variation of momentum ratio. Consequently, the mixing efficiency shows that maximum efficiency is at MR=3. after MR=3, mining efficiency decreases slightly.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.7
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• pp.544-551
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• 2014

A program of launch vehicle performance analysis is composed for the education of the conceptual design of launch vehicles and the requirement analysis for the propulsion system design. The program is applied for the mission analysis of space launch vehicles based on KSLV-II with liquid rocket boosters. The 75-ton class liquid rocket engine is assumed for the boosters by referring the mass ratio of KSLV-II second stage. The launch performance analysis is carried out for KSLV-II with 2, 3 and 4 boosters by targeting the circular orbit of 700 km altitude. The trajectory is assumed as two-dimension considering the variation of the flight environment. Payload of advanced KSLV-II could be increased to maximum 3 tons, though it is limited by the thrust performance of the upper stage.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.9
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• pp.32-37
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• 2003

The trend of miniaturization has been applied to the research on micro rockets resulting in prototype rockets fabricated by MEMS processes. In this paper, the development of three-dimensional micro rockets using micro milling as well as the results of combustion and flight tests are discussed. The body of rocket was made of 6061 aluminum cylinder. The three-dimensional micro nozzles were fabricated on brass by micro endmill with 127${\mu}m$ diameter. Two different micro nozzles were fabricated, one with 1.0mm of throat diameter and the other with 0.5mm. The total mass of rocket was 7.32g and that of propellant was 0.65g. The thrust-to-weight ratio was between 1.58 and 1.74, and the flight test with 45 degree launch angle form the ground resulted in 46m-53m of horizontal flight distance

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.4
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• pp.36-41
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• 2000

Mixing characteristics of unlike impinging streams double injectors have a great effect on combustion stability and combustion efficiency for liquid rocket engine. In this study a cold test was carried out, using water and TCE as simulants, in order to examine the effect of design parameters such as impingement angle, orifice diameter ratio and momentum ratio on the mass distribution and mixing quality.

• Journal of ILASS-Korea
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• v.25 no.3
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• pp.95-102
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• 2020

Six shear coaxial injectors for a 3 tonf-class liquid rocket engine using oxygen and methane as propellants were designed and manufactured by considering geometric design parameters such as a recess length and a taper angle. Cold-flow tests on the injectors were performed using water and air as simulants. By changing the water mass flow rate and air mass flow rate, the injection pressure drop under single-injection and bi-injection was measured. The discharge coefficients through the injector oxidizer-side and fuel-side were calculated and the discharge coefficient ratio between bi-injection and single-injection was obtained. Under single-injection, the recess served to reduce the injection pressure drop on the injector fuel-side. For the injectors without recess, the discharge coefficients under bi-injection were almost the same as those under single-injection. However, for the injectors with recess, the taper angle and bi-injection had a significant effect on the discharge coefficient.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.4
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• pp.96-103
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• 2005

The main purpose of launch vehicle is to insert satellite into a target orbit safely and correctly. To accomplish the main purpose of launch vehicle, the inserting velocity, inserting angle, and final mass of launch vehicle should be within the allowable range. In general, such requirements are satisfied with applying TCS(Thrust Control System) and TDS(Tank Depletion System), which manage thrust and mixture ratio by controlling propellant flow rate with thrust and mixture ratio control valves. In this study, the control characteristics of thrust and mixture ratio control valve were examined by PID control logic for stable operation of liquid-Propellant rocket engine at on-dosing point. The analysis on the control characteristics of control valves was done with AMESim code and the results from control valve test facility at KARI.

• The KSFM Journal of Fluid Machinery
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• v.5 no.4 s.17
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• pp.11-18
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• 2002

A turbopump system composed of two pumps and one turbine is considered. The turbine composed of a nozzle and a rotor is used to drive the pumps while gas passes through the nozzle and potential energy is converted to kinetic energy, which forces the rotor blades to spin. In this study, an aerodynamic design of turbine system is investigated with some pre-determined design requirements (i.e., pressure ratio, rotational speed, required power, etc.) following Liquid Rocket Engine (L.R.E.) system specifications. For simplicity of turbine system, impulse-type rotor blades for open-type L.R.E. have been chosen. Usually, the open-type turbine system requires low mass flow-rate compared to close-type system. In this study, a partial admission nozzle is adopted to maximize the efficiency of the open-type turbine system. A design methodology of turbine system was introduced. Especially, partial admission nozzle was designed by means of simple empirical correlations between efficiency and configuration of the nozzle. Finally, a turbine system design is presented for a 10 ton thrust level of L.R.E.