• 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 of Propulsion Engineers
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• v.10 no.3
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• pp.109-117
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• 2006

Hot firing performance test of hydrazine decomposition catalyst used for monopropellant thruster of satellite and launch vehicle was performed on the ground. A test equipment for hot firing performance measurement of catalyst test was developed in collaboration with Hanwha Corp., and the catalyst firing performance were tested with the equipment. After a reaction delay time, a catalyst activity and a granule stability were measured for 2 times, satisfactory results were obtained such as 25msec, 2%, $704^{\circ}C$ for each test items on the average. In addition, the current development status of domestic prototype catalyst and its decomposition performance test results are presented.

• 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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• 2012.05a
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• pp.167-171
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• 2012

A computational analysis of nozzle flow is conducted to investigate effects of the flight altitude on thrust performance. Reynolds-averaged Navier-Stokes equation with k-${\omega}$ SST(Shear Stress Transport) turbulence model is employed to simulate the nozzle flow in various altitude conditions, where continuum mechanics is to be valid. Thrust performance of the nozzle is exceedingly poor upto 10 km of flight altitude because of the irreversible phenomena such as shock and/or flow separation occurring inside the nozzle, whereas it is restored to the nominal value as the altitude is attained higher than 30 km.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.363-367
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• 2005

Latching valves are operated like solenoid valves by open/close command and they maintain final open or close commend without electric power source, so they are widely used in propulsion system of satellite and launch vehicle requiring reliability and being subject of restriction of power. This paper present design and test procedure of latching valve using permanent magnet polarized solenoid, which is developed for 45N Hydrazine propulsion system, to estimate feasibility of design and manufacture of latching valve.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.4
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• pp.1-8
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• 2014

Overexpanded flow of an axisymmetric thruster nozzle is numerically simulated to investigate effects of nozzle pressure ratio (NPR) on the shock structure and thrust performance. The Reynolds-averaged Navier-Stokes equations with k-${\omega}$ SST turbulence model are solved utilizing FLUENT solver. As the NPR is raised, thrust performance monotonically increases with the shock structure and flow-separation point being pushed toward the nozzle exit. It is also discussed that the flow structure at nozzle-exit plane which is immediately affected by a position of nozzle-interior shocks and expansion waves, has strong influence upon the thrust performance of thruster nozzle.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.5
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• pp.335-343
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• 2019

The third stage of the KSLV-II is equipped with the reaction control system that performs three axis-control during non-thrust coasting phase and performs a roll axis control during thrust phase. Toxic propellants such as hydrazine have been used for conventional rocket propulsions, however, recently, more studies have been conducted on the use of non-toxic eco-friendly propellants such as ADN and HAN. Especially, hydrogen peroxide has received a growing focus as an emerging propellant. It is considered an alternative of the toxic propellants because of economic advantage in producing the system, conducting operation test, and evaluation of the test result. In this paper, we describes the design, prototype, testing and evaluation of the test results with the 50 N-level hydrogen peroxide monopropellant thruster system which is currently under development.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.61-64
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• 2003

A comprehensive understanding is given for the hot-firing test results, which were obtained throughout the verification program of mono-propellant hydrazine thrusters producing 0.95 lbf (4.2 Newtons) of nominal steady-state thrust at an inlet pressure of 350 psia (2.41 ㎫). The scrutiny is made in terms of thrust and temperature behavior of steady state firing mode at the given propellant injection pressures of 400, 250, 100, and 50 psi. Engineering philosophy of data measurement and reduction is shortly mentioned, too.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.189-192
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• 2003

This paper summarizes a satellite program-specific performance requirements and test results for the verification of standard mono-propellant hydrazine thruster (MRE-1) producing 0.95 lbf (4.2 Newtons) nominal steady-state thrust at an inlet pressure of 350 psia (2.41 Mpa). Performance characteristics are shown in terms of thrust behavior at steady state and pulse mode firing. Hot firing test philosophy is briefly introduced, too.