• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.6
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• pp.581-586
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• 2008

The MEMS solid propellant thrusters have very low thrust level for applying to the propulsion system of micro/nano satellites or the side jet thruster of smart bombs. In this research, the fabrication possibility of planar type MEMS solid propellant thrusters that have enlarged burning surface area was examined and the safety of the structure of thruster during the firing test was confirmed. The performance of a micro igniter which is the key component of the MEMS solid propellant thruster was estimated by the ANSYS Icepak and evaluated by the experiment. Finally, the thrust was measured by the micro force sensor. The levels of thrust were 300, 600 mN in the case of K=15, 20.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.540-543
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• 2011

This work is devoted to an experimental investigation on conceptual design of dual consecutive stage micro plasma thruster (${\mu}PT$). Optimization study on the thruster configuration has been performed for various electrode gap distances from 1 mm to 2 mm and the hole diameter from 0.3 mm to 2 mm depending on desired operating conditions and corresponding nozzle design requirement. The operation of ${\mu}PT$ at low pressure from $10^{-1}$ Torr to $10^{-4}$ Torr and at various argon flow rates ranging from 5 sccm to 300 sccm has been studied to understand the physic of plasma and the gas dynamics in details. The specific impulse can reach up to 3000-4000 seconds at low power consumptions from 1 to 5 W. Image of exhaust plume from ${\mu}PT$ will be provided and electrical characteristics is also mentioned in this paper.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.628-633
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• 2004

초소형위성은 개발기간의 단축에 따른 발사 비용의 저감과 업무와 기능의 분산화에 따른 발사 실패의 리스크 저감 등 여러 장점이 있기 때문에 그 필요성이 높아지고 있다.(중략)

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.12
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• pp.1115-1120
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• 2007

This paper presents a novel mechanism of electro-spray micro-thruster featured by a nozzle with a conductive pole inside, referred to here as a pole type nozzle. And the effects of the pole type nozzle on the efficiency of the jetting are numerically and experimentally investigated. The electric voltage signal applied to the upper electrode plate, against the pole as the ground, allows a ejection of spray to take place. It is verified experimentally that the use of the pole type nozzle allows a stable and sustainable jetting mode of ejection for a wider range of applied voltages because it can concentrate the electric field more on the centre of the meniscus. According to results about size effect, experiments indicates that the proposed mechanism allows that operation of micro thruster at less than 500 volts through nanoscale nozzle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.593-596
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• 2004

We are developing a micro-solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft. The prototype has ø 0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 22 x 22 mm substrate. In previous studies, an impulse thrust of 4.6 x 10$^{-4}$ Ns was obtained in vacuum, but we found the problems of unacceptably low ignition success rate and incomplete combustion. This paper describes experiments to improve the ignition rate. In order to achieve this goal, we tried to solidify paste-like ignition aid (RK) on the ignition heaters with strong adhesion. To make the paste-like RK, isoamyl acetate was added to RK powder. We tested 9 rockets, but only 2 rockets were ignited with huge ignition energy. This is because the heat con-duction between the ignition heater and the RK was too low to ignite the RK, since dried RK had a lot of pores. Also, a large cavity was sometimes found just above the ignition heater.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.177-183
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• 2005

In this work, Inconel 600, Inconel 625 and Haynes 230 are welding using Nd:YAG laser and electron bear. These heat-resistant alloys are typically used for the liquid thruster. To examine, the affects of experimental parameters on weld ability, the plate welding of these materials were carried out using both Nd:YAG laser and electron bear. Also, the micro-structure, micro-hardness, and tensile strength of the specimens were analyzed. from the analysis of the experimental results of laser and electron beam welding, we have obtained the optimal welding conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.929-932
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• 2011

We developed a valve actuator controller for thruster system. This thruster system has four actuators and the actuator use a BLDC motor. Controller was made based on system and control requirement. The controller is consist of power, control and Amp. The control module use a micro-controller which is TMS320F28335 of Texas Instruments. It works for digital PID control and CAN communication and system control. The amp module for three phase BLDC motor use IGBT.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.91-94
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• 2007

The fabrication and firing test of the ignition system for a micro solid propellant thruster are described in the present paper. Pt igniter coil was patterned on the glass membrane that was fabricated by the wet etching process. The thickness of Pt layer was $2000{\AA}$ and the width of igniter pattern was $40{\mu}m$. The thickness and diameter of glass membrane were $15{\mu}m$ and 1 mm, respectively. Ignition test was performed. Successful ignition of HTPB/AP propellant was obtained with an ignition delay of 1.6 s at an input voltage of 12 V. The ignition energy was estimated to be 1.4 J.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.102-105
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• 2007

The performance evaluation of the micro thruster utilizing hydrogen peroxide decomposition is described. The catalyst bed was made of porous ceramic material($Isolite^{(R)}$) with large surface to mass ratio. 14%wt platinum was loaded on the catalyst support as a catalyst. Hydrogen peroxide with 85% concentration was used as a monopropellant. The length of the catalyst bed and the feed pressure of the hydrogen peroxide were taken as the parameters for the experiment. All experiments were carried out under cold start condition for 30 seconds. The $c^*$ efficiency was evaluated for each test case using measured pressure data. For the catalyst support length of 30 mm and feed pressure at 5.51 bar, satisfactory $c^*$ efficiency beyond 95% was observed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.109-112
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• 2008

A performance of micro thruster was measured with catalyst bed that was prepared by different catalyst loading condition for the optimization of catalyst bed size. Among the catalyst loading conditions, pH level of precursor solution was changed by several solutions like Nitric acid or Sodium hydroxide. For the each case, it was heated at different drying temperatures that can affect the phase of catalyst loaded on support. From these results, it was studied that the effect of catalyst loading condition on the performance. 90wt% hydrogen peroxide was used as a monopropellant, and platinum was chosen as a catalyst. Characteristic velocity efficiency and temperature efficiency were used for the performance evaluation.