• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.221-224
• /
• 2010

Recently, due to the enormous cost for sending a satellite into an orbit, small and more reliable satellites have been more demanded. An introduction of new binders(HTPB, GAP) and new oxidizers made great improvements of the large thrust modulation. In order to make cost reduction, one prefers to the low melting temperature thermoplastic propellant reforming the manufacturing process dramatically. Solid propellant rockets have been had a problem of the injection accuracy into orbit, but PBS(Post Boost Stage) using a liquid mono-propellant improves the injection accuracy. This paper also gives the direction of the advanced nozzle materials and the motor case.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.11a
• /
• pp.203-206
• /
• 2008

Kick motor(KM) for KSLV-I second stage propulsion system is the main hardware that is necessary for launching satellite to it's track. The mass of the kick motor changes with combustion time because the heat insulator is ablated and propellant is used and slag is piled up. We predicted mass change with the flight time using ground combustion data of KM composed of case, propellant, nozzle, ignitor and slag. The mass prediction of kick motor can be used for calculating the two stage mass and center of gravity history.

• Aerospace Engineering and Technology
• /
• v.2 no.2
• /
• pp.76-82
• /
• 2003

To guarantee the proper functions of a satellite in the extreme space environment, the several test models are developed generally. There are advantages that the design and the analysis of Flight Model(FM) can be validated through these test models, and the functional reliabilities can be increased by reflecting the modifications on the final design of FM. The integration and test of Structure & Thermal Model(STM) of KOMPSAT, being currently developed, have been completed. In this paper, the processes of design and analysis of the STM propulsion system, one of the KOMPSAT modules, are described.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.11a
• /
• pp.827-829
• /
• 2011

Vacuum facility is required for high altitude space environment test to develop satellites or space launch vehicles. We, at Chungnam, National University, developed turbo molecular pump vacuum test facility up to $1.0{\times}10-6$ torr to simulate 200 km altitude environment. In this paper, we present some preliminary vacuum performance test results.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.251-254
• /
• 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
• /
• 2003.05a
• /
• pp.189-192
• /
• 2003

This paper summarizes a satellite program-specific performance requirements and test results for the verification of standard mono-propellant hydrazine thruster (MRE-1) producing 0.95 lbf (4.2 Newtons) nominal steady-state thrust at an inlet pressure of 350 psia (2.41 Mpa). Performance characteristics are shown in terms of thrust behavior at steady state and pulse mode firing. Hot firing test philosophy is briefly introduced, too.

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

• Journal of the Korean Society of Propulsion Engineers
• /
• v.5 no.4
• /
• pp.67-74
• /
• 2001

Based on the national space development project, the necessity of developing liquid propellant rocket engine is revealed to secure the basic technology for the development of individual artificial-satellite launcher. Consequently, KARI (Korea Aerospace Research Institute) is developing a liquid propellant rocket engine for the KSR-III. Currently, a prototype engine using kerosene/LOx which produces 13-ton thrust is designed, fabricated and tested. In this paper, test procedure and technique for liquid propellant rocket engine are introduced with the analysis of static and dynamic test data.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 1997.11a
• /
• pp.21-21
• /
• 1997

인공위성용 추진제 탱크를 개발하기 위해 여러 설계인자를 설정하여 각 인자가 탱크벽면에 미치는 응력분포 영향을 구하고, 또한 최적의 인자 값을 구하기 위해 각 인자의 변화에 따라서 구조해석을 수행하였다. 탱크 지지부 위치와 탱크 벽면 두께 변화에 따른 탱크 벽면에 미치는 응력분포 영향을 고찰하기 위해 1/4 모델을 설정하였고, 연료배출구의 위치변화(경사각돈)에 따른 응력분포는 1/2 모델을 설정하여 해석을 하였다. 탱크에 작용하는 하중은 연료압력에 의해 발생하는 정하중(350 psi)을 가하며 또한, 발사 시 발사체로부터 전달되는 최대동하중(llg)을 고려하였다. 그리고, 탱크가 인공위성에 장착될 때에 발생하는 다양한 장착조건에 대해서 구조해석을 수행하였고, 추진제 배출구 각도가 $0^{\circ}$ 에서 $25^{\circ}C$까지 변화할 때 탱크 벽면에 미치는 응력분포 영향을 구했다. 그래서 각 조건에서 구한 상당응력분포와 인자의 최적 값은 추진제 탱크를 설계하기 위한 기초적인 자료로 활용하고자 한다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.314-317
• /
• 2004

On 9/5/2003, the planet probe “HAYABUSA” as MUSES-C project was launched by The Institute of Space and Astronautical Science. “HAYABUSA” has microwave discharge ion engines and these engines are characterized by their high efficiency and specific impulse in comparison with chemical engine. A large ion engine can be used as a planet explorer, while a small ion engine can be used as attitude control of small satellite. We have been developing a high thrust density microwave discharge ion engine using “Multi-Monopole Antenna”. The performance of this engine are: ion cost of 344W/A, propellant utilization efficiency of 52% and thrust density of 0.055mN/$\textrm{cm}^2$ for Kr gas flow rate of 2.5sccm, microwave(2.45㎓) power of 32W and acceleration voltage of l.4㎸.