• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.31 no.2
• /
• pp.89-95
• /
• 2003

Propulsion system of the current developing satellite is roughly composed of propellant tank and four major modules. Each module prevides the pulse momentum for spacecraft attitude control, filling/draining of propellant and pressurant, propellant filtering, and the change of flow passage in the spacecraft emergency situation, respectively. These modules will be fixed on the propulsion platform with their suitable mounting brackers, so the brackets shall be designed sufficiently to support a function of the modules under launch environment and on-orbit condition. The purpose of this article is to check if all the bracket designs satisfy the defined structural requirements through finite element analysis, and then to verify structural safety.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.365-368
• /
• 2009

A new concept of a small thruster for altitude control of a micro/nano class satellite is developed, which utilizes the pulsed laser energy. As the laser-based thruster does not require burning of any fuel, it gives promise of small satellite design criteria, namely light weight and cost effectiveness. In this paper, we develop gel-type material for generating strong plasma plume for enhancing thrust for propulsion. Moreover, we quantify the level of thrust via the momentum coupling coefficient measured by the pendulum system. We discover that the driving force is significantly improved via the gel-typed propellant for laser ablation.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.10 no.3
• /
• pp.26-31
• /
• 2006

Through the KOMPSAT program, Koreanization of thruster have been carried out successfully, but there are still many difficulties in Koreanization of most core parts of propulsion system. Because the development of core parts is essential to participate in the advanced nations, KARI has carried out development of Fill/Drain valve for propellant/pressurant supply of satellite, which has high possibilities to be koreanized, with Hanwha Corp.. This paper summarizes overall processes of development including design, manufacturing and test, and finally 4 sets of modules were successfully made. Also the satisfaction of performance requirements are verified through performance tests.

• Proceedings of the KSRS Conference
• /
• v.1
• /
• pp.38-38
• /
• 2006

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.838-842
• /
• 2017

The thruster valve, which is one of the key components applied to the mono-propellant system for the satellite, has a circular plate spring structure. It can be designed as a structure that does not have positional deformation and particles by friction and repetitive motion. In this study, finite element analysis and verification were performed by setting the width of the circular plate spring as a design parameter with the material, thickness and radius of the circular plate spring as fixed variables. The linearity of the spring constant is shown by the graph that is spring force with displacement. It is confirmed that the optimization design of the circular plate spring is possible by the spring force tendency according to the total area of circular plate spring.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.35 no.4
• /
• pp.323-328
• /
• 2007

A direct application of liquefied gas propellants to a typical small satellite cold gas propulsion system was analyzed. Performance of systems using liquefied gas propellant under consideration was compared to that of a nitrogen cold gas propulsion system. Liquefied gas propellant propulsion system's performance, required tank volume, and required propulsion system mass has been calculated at the same mass, volume, and total impulse condition of a typical nitrogen cold gas propulsion system. It was found that the liquefied gas propulsion system has advantages in performance, volume, and mass, compared to a typical nitrogen cold gas system, and can be directly applied to a cold gas propulsion system.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.42 no.12
• /
• pp.1051-1056
• /
• 2014

A Fuel Feeding Unit of electric propulsion system has been developed for the small-satellite applications. The fuel feeding unit stores the xenon gas with high pressure and density as a fuel. Xenon can affect to system stability since xenon has the transient condition under the critical point which is in ambient temperature on the launch environment. The functional and structural stability on the launch environment needs verification through the ground tests. The design points and verification tests of the system were discussed and test results were described on this text. The system stability on the launch environment was verified through these verification tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.03a
• /
• pp.283-292
• /
• 2008

The foremost criterion in the design of a Satellite Launch Vehicle(SLV) is its performance capability to boost the designated payload to the desired mission orbit; it starts from focusing on the SLV configuration to achieve the velocity requirements($}\Delta}V$) for the mission. In this paper we review an analytical approach which is suitable enough for preliminary conceptual design and is used previously to optimize stage configurations for Two Stage to Orbit SLV for Low Earth Orbit(LEO) Missions; we have extended this approach to Three Stage to Orbit SLV and compared different propellant options for the mission. The objective is to minimize the Gross Lift off Weight(GLOW). The primary performance figures of merit were the total inert weight of the SLV and the payload weight that the SLV could lift into LEO, given candidate propulsion systems. The optimization is achieved by configuring the $}\Delta}V$ between stages. A comparison of configurations of single-stage and multi-stage SLVs is made for different propellants. Based upon the optimized stage configurations a comparative performance analysis is made between Liquid and Solid fueled SLV. A 3 degree of freedom trajectory-analysis program is modeled in SIMULINK and used to conduct the performance analysis. Furthermore, a cost analysis is performed on our stage optimized SLVs. The cost estimation relationships(CER) used give us a comparison of development and fabrication costs for the Liquid vs. Solid fueled SLV in man years. The pros and cons of the production, operation ability, performance, responsiveness, logistics, price, shelf life, storage etc of both Solid and Liquid fueled SLVs are discussed. The statistics and data are used from existing or historical(real) SLV stages.

• Journal of Astronomy and Space Sciences
• /
• v.26 no.3
• /
• pp.403-416
• /
• 2009

COMS (Communication, Ocean and Meteorological Satellite) is a geostationary satellite and developed by KARl for communication, ocean and meteorological observations. It will be tested under vacuum and very low temperature conditions in order to verify thermal design of COMS. The test will be performed by using KARI large thermal vacuum chamber, which was developed by KARI, and the COMS will be the first flight satellite tested in this chamber. The purposes of thermal balance test are to correlate analytical model used for design evaluation and predicting temperatures, and to verify and adjust thermal control concept. KARI has plan to use heating plates to simulate space hot condition especially for radiator panels of satellite such as north and south panels. They will be controlled from 90 K to 273 K by circulating GN2 and LN2 alternatively according to the test phases, while the main shroud of the vacuum chamber will be under constant temperature, 90 K, during all thermal balance test. This paper presents thermal modelling including test chamber, heating plates and the satellite without solar array wing and Ka-band reflectors and discusses temperature prediction during thermal balance test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.597-600
• /
• 2004

Fabrication technique and performance test of catalytic micro propulsion are treated based on MEMS technology. This propulsion is designed to use hydrogen peroxide as liquid mono-propellant for attitude control of pica-satellite. The propellant is fed into the micro reactor channel and decomposed into hot gas yielding controllable thrust by catalyst. In order to increase the efficiency of the reaction that depends on the contact area of propellant and catalyst, porous surface formation on the channel accompanied by platinum particle deposition has been performed using H$_2$PtCl$_{6}$ solution as a precursor. Several thrusters were fabricated in different concentration of H$_2$PtCl$_{6}$ solution to determine the best quantity of Pt particles. For the comparison of the performance of each thruster, the volume of oxygen generated by the decomposition of hydrogen peroxide and the thrust were measured.red.