• International Journal of Aeronautical and Space Sciences
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• v.17 no.4
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• pp.526-534
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• 2016

A MEMS solid propellant thruster array shall be operated within an allowable range of operating temperatures to avoid ignition failure by incomplete combustion due to a time delay in ignition. The structural safety of the MEMS thruster array under severe on-orbit thermal conditions can also be guaranteed by a suitable thermal control. In this study, we propose a thermal control strategy to perform on-orbit verification of a MEMS thruster module, which is expected to be the primary payload of the STEP Cube Lab mission. The strategy involves, the use of micro-igniters as heaters and temperature sensors for active thermal control because an additional heater cannot be implemented in the current design. In addition, we made efforts to reduce the launch loads transmitted to the MEMS thruster module at the system level structural design. The effectiveness of the proposed thermo-mechanical design strategy has been demonstrated by numerical analysis.

• 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.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.2
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• pp.80-85
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• 2014

Side jet thruster using nozzle closure separation device provides a solid rocket with a trajectory shift function. Side jet thruster consists of low combustion temperature propellant, neutral type propellant grain and nozzle closure separation device. If a trajectory shift is required, side jet thrust is generated on the rocket by separating some nozzle closures located in the opposite direction to thrust. After completing trajectory shift, the other nozzle closures located in the thrust direction are separated to cease side jet thrust. The operation process is verified through ground static test. The result in this study can be applied to changing rocket trajectory by controlling side jet thrust through nozzle closure separation.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.364-367
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• 2011

The development of the DCS thrust unit during the attitude control thruster of the launch vehicle and guided missile is introduced. The DCS thrust unit using solid propellants based on a two-axis control is designed and through the thermo-structural and flow analysis is designed in detail. The performance of the thrust unit based on the detail design is demonstrated through a combustion test.

• 제어로봇시스템학회:학술대회논문집
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• 2005.06a
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• pp.352-355
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• 2005

This paper describes an engineering model development of a pulsed plasma thruster, which is capable of an impulse bit of 20uNs and a specific impulse of 800s. The solid fuel which is Teflon allows for a self-contained, inert and stable propellant system. And, the PPT technology makes it possible to consider a revolutionary attitude control system (ACS) concept providing stabilization and pointing accuracies previously obtainable only with reaction wheels, with reduced mass and power requirements.

• 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.

• Journal of the Computational Structural Engineering Institute of Korea
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• v.24 no.6
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• pp.699-705
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• 2011

High-maneuver missile is a projectile which demands a strong momentum at short time. To produce a necessary thrust for the flight, the gas of high temperature and pressure is generated through explosive combustion of solid propellant, and a great thrust can be obtained by expanding this high temperature and pressure gas. Although the operating time of a rocket motor is less than a few seconds, a failure of part or ablation near the throat of nozzle may take place during the expansion of high temperature and pressure gas for great thrust. In other words, for the precise control of a missile an exact stress analysis considering both, the thermal stress caused by the heat transfer between combustion gas and wall, and the mechanical stress caused by the pressure change in the flow, should be considered first. In this connection, this study investigated the safety, as a point of view of stress and melting point of the material, of the pre-designed thrust generating structure which is subjected to high temperature and pressure as a function of motor operating time.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.2
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• pp.156-164
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• 2016

Qualification model(QM) of main payloads including concentrating photovoltaic system using fresnel lens, heating wire cutting type shockless holding and release mechanism, and MEMS-based solid propellant thruster have been developed for the STEP Cube Lab.(Cube Laboratory for Space Technology Experimental Project), which is a pico-class satellite for verification of core space technologies. In this study, we have verified structural safety and functionality of the developed payloads under a qualification temperature range through the QM thermal vacuum test. Additionally, a reliability of thermal model of the payloads has been confirmed by performing a thermal correlation based on the thermal balance test results.