• Journal of Aerospace System Engineering
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• v.9 no.2
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• pp.13-18
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• 2015

Micro planar type liquid propellant thruster was fabricated by MEMS manufacturing process for micro/nano satellites applications. 90 wt.% hydrogen peroxide was used as propellant and for propellant decomposition, Pt/Al2O3 was used as catalyst. Micro thruster structure was made by 5 photosensitive glasses patterned with thruster component profiles. Objective thrust was 50 mN and required hydrogen peroxide mass flow was 2.1 ml/min, which was supplied by syringe pump and teflon tube in experimental test. Performance test said that average steady thrust was approximately 30 mN, around 60% of objective thrust, and transient time was about 5 sec. It is estimated that extended response time was due to high thermal energy loss of micro scale thruster and low enthalpy input by propellant mass flow.

• 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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• 2008.05a
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• pp.111-114
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• 2008

The performance evaluation of a micro monopropellant thruster is described in this paper. The platinum which has stable physical and chemical properties is used as a catalyst. The ceramic form, named Isolite is chosen as a catalyst bed. Hydrogen peroxide whose concentration is up to 90wt% is supplied into thruster by direct pressurization with nitrogen gas. The $c^*$ efficient is calculated to evaluate the performance of a thruster. The pressure rise time, pressure decent time and reaction delay are synthetically considered.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.151-154
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• 2007

The purpose of this study is to identify the basic concept of thruster design through the visualization firing test on a hydrazine thruster. We designed the visual catalyst bed on the basis of the 1lbf hydrazine thruster for a low earth orbit satellite and observed visually the internal catalyst bed reaction.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.137-141
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• 2004

Many monopropellant thrusters use a catalyst for decompose the propellant, hydrazine. The catalyst directly affects the thruster performances and lifetime. Therefore, it is important to confirm that the catalyst is suitable for our thrusters. Until 2002, we used She1l405 catalyst, for satellite RCS thrusters, and H-IIA and M-V launch vehicle upper-stage RCS thrusters. In 2002, however, Shell Chemical Inc. ceased manufacturing She1l405 catalyst and transferred the product to AEROJET, where it was renamed S405. We subsequently investigated the characteristics of AEROJET's S405 catalyst and SOLVAY's KC12GA catalyst, (SOLVAY is a Belgian chemical company, and KC12GA is used for ASTRIUM's thruster) and conducted thruster firing tests using the new catalysts. After conducting, we confirm that the KC12GA catalyst was suitable for our thrusters, and decided to use KC12GA for two satellite programs.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.378-382
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• 2008

A 50 N monopropellant thruster being developed for attitude control in a variety of aerospace application systems is described in this paper. Ninety percent hydrogen peroxide was selected as a propellant, since it is much less hazardous than hydrazine. A scaled down thruster with aluminum oxide loaded with the platinum in the reaction chamber was tested to determine propellant decomposition onto a catalyst. A scaled up 50 N thruster, with a catalyst bed of 3 cm in diameter and 4 cm in length, was evaluated by decomposition efficiency based on temperature, ${\eta}_T$, efficiency of characteristic velocity, ${\eta}_{C^*}$, and measurement of thrust. The performance of a 50 N thruster was 40.5 Newton in thrust, about 100 % in ${\eta}_T$, and 98 % in ${\eta}_{C^*}$, and 125 sec in specific impulse at sea level.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.5
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• pp.496-501
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• 2009

The thruster head unit of a 300 W cylindrical Hall thruster was developed for the propulsion system of small satellites. The magnetic topology in the thruster channel is a key parameter to achieve high performances. Two types of magnetic circuit structures were designed and manufactured to compare the thrust levels and efficiencies. Also the endurance test was conducted to measure the stable operation duration of the thruster head and to find degree of erosion after extended operation.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.77-84
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• 2006

With the development of micro/nano spacecraft, concepts of micro propulsion are introduced for orbit transfer and drag compensation as well as attitude control. Micro solid propellant thruster has been attention as one of possible solution for micro thruster. In this paper, micro solid propellant thruster is introduced and research on basic components of a micro solid propellant thruster is reported. Micro Pt igniter was fabricated through negative patterning and quantitative effect of geometry was estimated. The characteristic of HTPB/AP solid propellant was investigated to measure the homing velocity. A combustion chamber was fabricated by means of anisotropic etching of photosensitive glass. Finally, micro solid propellant thrusters having various geometries were fabricated and tested.