• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.5
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• pp.432-439
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• 2015

MEMS solid thruster module is composed of solid thruster and its control board. It was developed for the purpose of an academic research. Therefore, thermo-mechanical design and verification for space usage were not considered in the design phase. To mount it on a cube satellite without any design modification, technical efforts at the system level structure design is required. In this study, we proposed a structural design concept to mount the MEMS thruster module by using brackets for guaranteeing structure safety under launch loads and easier mating and de-mating of MEMS thruster module during test phase. The effectiveness of the design has been verified through structural analysis and vibration test. In addition, electrical connection method using spring pins between MEMS thruster and control board is effective for guaranteeing the structural safety under launch vibration loads.

• 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.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.9
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• pp.802-808
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• 2014

MEMS based solid propellant thruster researched for the purpose of an academic research will be verified at space environment through CubeSat program. For this, the temperature of the MEMS thruster should be within allowable operating temperature range by proper thermal control to prevent the ignition failure caused by ignition time delay and to guarantee the structural safety of the MEMS thruster in the low temperature. In this study, we proposed an effective thermal control strategy, that is to use micro-igniter as a heater and temperature sensor for active thermal control instead of using additional heater. The effectiveness of the strategy has been verified through on-orbit thermal analysis of CubeSats with MEMS thruster.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.1
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• pp.1-8
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• 2015

Micro igniter on the glass membrane for MEMS thruster was developed. The stability of the micro igniter by using a glass membrane with a thickness of tens of microns was improved. The micro igniter was fabricated by anisotropic wet etching of photosensitive glass and deposition of Pt/Ti for electric heat coil. The solid propellant was loaded into the propellant chamber without an especial technique due to the high structural stability of the glass membrane. Ignition tests were performed successfully. The minimum ignition delay was 27.5 ms with an ignition energy of 19.3 mJ.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.225-228
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• 2007

Micro solid propellant thruster is the most feasible for development with current MEMS. Basic components of micro solid propellant thruster are diverging nozzle, micro igniter, combustion chamber, and solid propellant. Micro nozzles and micro chambers were fabricated using photosensitive glass by anisotropic wet etching technique. Micro Pt heaters on glass membrane which ignited solid propellant were developed. Components of thruster were integrated. Successful ignition was observed.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.6
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• pp.1-7
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• 2017

The performance of a MEMS solid propellant thruster was predicted and analyzed through internal ballistics model and CFD analysis. The nozzle throat was $416{\mu}m$, and the area ratio of the nozzle was 1.85. As a result of the internal ballistics model, chamber pressure increased up to 197 bar and the maximum thrust was 3,836 mN. In CFD analysis, the chamber pressure of the internal ballistics model was applied as the operating pressure, and the CFD model was divided into an adiabatic and a heat loss model. As a result, the maximum thrust of the adiabatic model was 14.92% lower than that of the internal ballistics model, and the effect of heat loss was insignificant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.757-763
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• 2017

The combustion test of LMP-103S, a propellant based on ADN(Ammonium Dinitramide), was performed with a 50 mN scale micro-thruster. The micro-thruster was made with photosensitive glass using MEMS manufacturing process. $Pt/{\gamma}-Al_2O_3$ was used as a catalyst to decompose LMP-103S. After injecting 90 wt.% hydrogen peroxide into combustion chamber to preheat the catalyst, LMP-103S was injected for the combustion test. As a result, the ignition and combustion of LMP-103S was confirmed in platinum catalyst environment with the combustion chamber temperature going up to $650^{\circ}C$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.251-254
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• 2006

Micro thruster is a key technology in the micro/nano satellite. MSPT has been attracted attention as a one of possible solution for micro thruster MSPT as a systems four components. It is composed of nozzle, igniter, combustion chamber and propellant. This paper surveys varioud MSPTs which have been reported. The model of MSPT arrays for total impulse of 1 mNs is proposed. Combustion chamber is designed and fabricated.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.6
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• pp.111-117
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• 2018

In this study, an MEMS thrust measurement system was designed and a study on the performance prediction of system was performed to evaluate the performance of micro thruster. Thrust measurement system consists of beam, membrane, and piezoresistive sensor. An FEM analysis was carried out to verify the stability of the system, confirm the stress variation at the beam, and position the piezoresistive sensor. The stability of the designed system was verified by comparing the yield strength of the material with the maximum stress. The piezoresistive sensor was designed to be 20% of the length of the beam to obtain a high gauge factor. The size of the membrane and the beam of the reference model were designed to be $15mm{\times}15mm$, and $500{\mu}m{\times}500{\mu}m$, respectively.