• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.42-45
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• 2012

As a basic research for the development of Korean lunar lander, propulsion system development for ground test is in progress. Design target is 220 N in ground thrust at 130 g/s flow rate, 200 psi chamber pressure. For the performance improvement, two type injector and catalyst bed was designed. For ground test, thrust measurement system using LM guide was developed and test was performed. The result shows 214.1 N thrust in atmosphere condition at 126.6 g/s flow rate.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.135-138
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• 2011

As a basic research for the development of Korean lunar lander, propulsion system development for ground test is in progress. Thrust for descent is 200 N class. Design target is 220 N in vacuum thrust at 100 g/s flow rate, 200 psi chamber pressure. For ground test, thrust measurement system using LM guide was developed and test was performed. The result shows 160 N thrust in atmosphere condition at 210 psi chamber pressure.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.9
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• pp.649-657
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• 2019

Hydrazine, which is used as a representative monopropellant, is an extremely poisonous substance and has a disadvantage that it is harmful to the human body and is very difficult to handle. In recent years, research on the development of non-toxic and environmentally friendly propellants has attracted much attention. Ammonium dinitramide(ADN) based propellant developed by Swedish Space Corporation has superior performance to hydrazine and has been commercialized through performance verification in space environment. On the other hand, the exhaust gas from a thruster nozzle collides with a satellite while it is spreading in the vacuum space, thermal load and surface contamination may occur and may reduce the performance and lifetime of the satellite. However, a study on the effect of the exhaust gas of the green propellant thruster on the satellite has not been conducted in earnest yet. Therefore, the exhaust gas behavior in space was analyzed in this study for the ADN based green monopropellant using Navier-Stokes equations and the DSMC method. As a result, it can be expected to be used as design validation data in the development of satellite when using the ADN based green monopropellant.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.1-8
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• 2013

A liquid-monopropellant hydrazine thruster has several outstanding advantages such as relatively-simple structure, long/stable propellant storability, clean exhaust products, and so on. Therefore hydrazine thruster has such a wide application as orbit and attitude control system (ACS) for space vehicles. A hydrazine thruster with the medium-level thrust to be used in the ACS of space launch vehicles (SLV) has been developed, and its ground firing test result is presented in terms of thrust, impulse bit, temperature, and chamber pressure. It is verified through the performance test that the response and repeatability of thrust are very excellent, and the thrust efficiencies compared to its ideal requirement are larger than 93%.

• 한국연소학회:학술대회논문집
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• 2002.11a
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• pp.237-240
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• 2002

Micro catalytic reactors are alternative propulsion device that can be used on a nano satellite. When used with a monopropellant, $H_2O_2$, a micro catalytic reactor needs only one supply system as the monopropellant reacts spontaneously on contact with catalyst and releases heat without external ignition, while separate supply lines for fuel and oxidizer are needed for a bipropellant rocket engine. Additionally, $H_2O_2$ is in liquid phase at room temperature, eliminating the burden of storage for gaseous fuel and carburetion of liquid fuel. In order to design a micro catalytic reactor, an appropriate catalyst material must be selected. Considering the safety concern in handling the monopropellants and reaction performance of catalyst, we selected hydrogen peroxide at volume concentration of 70% and perovskite redox catalyst of lantanium cobaltate doped with strondium. Perovskite catalysts are known to have superior reactivity in reduction-oxidation chemical processes. In particular, lantanium cobaltate has better performance in chemical reactions involving oxygen atom exchange than other perovskite materials. In the present study, a process to prepare perovskite type catalyst, $La_{0.8}Sr_{0.2}CoO_3$, and measurement of its propellant decomposition performance in a test reactor are described.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.6
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• pp.19-26
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• 2014

A ground hot-firing test (HFT) was carried out to make a close examination into the pressure instability for the 70 N-class hydrazine thruster under development. Monopropellant grade hydrazine was adopted as a propellant for the HFT, and catalyst-bed was filled with $Ir/Al_2O_3$ catalyst. In order to investigate the effects of thrust-chamber diameter on combustion stability, evaluation tests for the development models were performed on three kinds of lower thrust chambers having the length-to-diameter ratio (L/D) of 1.03, 1.13, and 1.26. As results, it was found that low frequency instability (~ 50 Hz) was inherent in the models, and in addition, increase of the L/D and decrease of the operating pressure led to an amplification of pressure oscillation in the test condition specified.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.110-113
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• 2007

For the development of hydrazine decomposition catalyst, Hot-fire test to verify performance of catalyst is required. The purpose of a long-life firing test is to demonstrate the capability of a design to perform for the maximum duration or cycles of operation. This paper describes the progress of the catalyst performance test, explains the test matrix, and presents the test results.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.27 no.10
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• pp.1377-1384
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• 2003

An analysis has been performed for active thermal control of the KOMPSAT monopropellant rocket engine assembly, i.e., dual thruster module(DTM). The main efforts of this work have been directed at determining proper heater sizes for propellant valves and catalyst beds necessary to maintain their temperatures within specified temperature ranges under KOMPSAT environment and operational conditions. The TAS incorporated with TRASYS thermal radiation analyzer was used to establish a complete heat transfer model which allows to predict the DTM temperature as a function of time. The thermal analysis has been performed in transient mode to verify the appropriate power for catalyst bed heaters necessary to increase catalyst bed temperature to the required value within a specified period of time. Similar analysis has been executed to validate the heater power for the thermostatically controlled primary and redundant heater circuits used to prevent hydrazine freezing, i.e., single fault. Moreover the effect of the radiative property of thermal control coating of heat shield was examined. Thruster firing condition was also simulated for the heat soakback condition. As a consequence, all thermal analysis results for DTM satisfactorily met the thermal requirements for the KOMPSAT DTM under the worst case average voltage, i.e. 25 volt.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.41 no.4
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• pp.299-304
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• 2013

The HAN which is an ionic liquid is a non-toxic monopropellant with high storability, and its specific impulse can be increased by blending methanol, thereby it can substitute the hydrazine. The HAN was synthesized by acid-base reaction of hydroxylamine and nitric acid, and the blending ratio of HAN and methanol is 8.2:1. The iridium catalyst was used to decompose the HAN, and 1 N class thruster with shower head type injector having one orifice was used to evaluate the HAN/Methanol propellant. The thermal stability of distributor was increased by using ceramic material to endure the high temperature of product gas. The preheating temperature of catalyst should be $400^{\circ}C$ at least for the complete decomposition. The feeding pressure should be increased to increase the $C^*$ efficiency, thereby the decomposition performance was decreased upstream catalyst, and the performance of thruster was decreased. The fine metal mesh was inserted after the injector to improve the atomization of propellant, thereby it can settle the performance decrease problem. The phenomenon of performance decrease was remarkably improved owing to the insertion of fine metal mesh.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.1
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• pp.50-57
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• 2014

A ground hot-firing test (HFT) was conducted to take out the optimal design configurations for the thrust chamber of 70 N-class liquid rocket engine under development. Monopropellant grade (purity: ${\geq}98.5%$) hydrazine was adopted as a propellant for the HFT, and three kinds of thrust chambers having characteristic lengths ($L^*$) of 2.79, 2.95, and 3.13 m were selected for their performance evaluation. It is revealed through the test and evaluation that the increase of the $L^*$ leads to a performance degradation in the test condition specified, and pulse response performance of the development model shows superior characteristics to commercialized hydrazine thrusters.