• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.336-339
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• 2011

Hot-fire test facility is one of the most important infrastructure for thruster development and evaluation. During the past three years, Korea Aerospace Research Institute (KARI) and Hanwha Corporation have successfully performed the construction of hot-fire test facility for medium-scale monopellant thruster to the maximum 200N thrust level. In general, thruster hot-firing test should be performed in vacuum conditions to simulate space environment. The hot-fire test facility is divided into three subsystems, vacuum system, propellant supply system and data measurement & control system. The goal of this facility is to extend the capability from small thruster for satellite mission to medium-scale thruster for launch vehicle and lunar mission. In this paper, the progress and overview for thruster hot-fire test facility was introduced and test results were also presented.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.6
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• pp.103-120
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• 2010

A development history of the thrusters used for space-vehicle orbit maneuver/attitude control is reviewed with their performance characteristics. Especially, a scrutiny is made for the current and practical application of TVC/Gimbal/Thrusters to the roll/pitch/yaw-axis control of each stage of launch vehicles. It is well perceived that a precise 3-axis attitude control system (ACS) must be equipped on the final stage of space launch vehicles (SLV) for an attainment of orbit-insertion accuracy. Under the superior reliability as well as moderate performance features, the monopropellant hydrazine thrusters occupy most of the SLV's 3-axis ACS currently operated. Domestic development status of the medium-thrust-level thruster is shortly introduced, finally.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.6
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• pp.68-74
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• 2014

During the life firing test of 1 N-class thruster development model, pulse mode performance and performance changes were examined. The deviation of pulse mode response time according to thruster feed pressure was relatively small and the resultant ignition delay, response time, tail-off time were 32-35 ms, 86-91 ms, 89-98 ms, respectively. For the stabilized pulse region the impulse bit revealed the outstanding reproducibility of 1.41, 1.32, 2.10% at $3{\sigma}$. During the life firing test, the impulse bit was decreased with limited amounts, therefore the pulse mode performance could be considered to be maintained. The thrust centroid was also maintained during the life firing test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.746-752
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• 2017

High test hydrogen peroxide has been widely developed as green propellant for thrusters. Hydrogen peroxide is decomposed in the catalyst bed to produce the thrust. Catalyst bed design optimization is considered through existing model for catalyst beds. To verify the model, static firing tests were conducted under various conditions using a 100 N scale $H_2O_2$ monopropellant thruster. Temperature and pressure estimations from the model were well correlated to the experimental data. The model is used to obtain optimal design parameters by analyzing the catalyst capacity and pressure drop data for various simulated conditions. Catalyst beds can be optimized from the analysis of the catalyst capacity and pressure drop correlation through catalyst bed modeling.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.6
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• pp.19-26
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• 2014

A ground hot-firing test (HFT) was carried out to make a close examination into the pressure instability for the 70 N-class hydrazine thruster under development. Monopropellant grade hydrazine was adopted as a propellant for the HFT, and catalyst-bed was filled with $Ir/Al_2O_3$ catalyst. In order to investigate the effects of thrust-chamber diameter on combustion stability, evaluation tests for the development models were performed on three kinds of lower thrust chambers having the length-to-diameter ratio (L/D) of 1.03, 1.13, and 1.26. As results, it was found that low frequency instability (~ 50 Hz) was inherent in the models, and in addition, increase of the L/D and decrease of the operating pressure led to an amplification of pressure oscillation in the test condition specified.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.6
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• pp.30-37
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• 2013

A ground hot-firing test was conducted to take out the optimal design configurations for the catalyst bed of liquid-monopropellant hydrazine thruster which could be used for primary engine or attitude control thruster of space vehicles. Performance characteristics with the variation of thrust-chamber length are investigated in terms of thrust, specific impulse, chamber pressure, characteristic velocity, and hydrazine decomposition rate. Additionally, the correlations between propellant-supply pressure and performance parameters are given. As results, increase of catalyst-bed length leads to performance degradation in this test condition, and also decreases propellant consumption efficiency with the supply pressure variation.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.8
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• pp.747-752
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• 2007

Domestic satellite development programme is generally classified into two categories: COMS as GEO satellite and KOMPSAT as LEO one. Each satellite has the on-board propulsion system fulfilling its own mission requirements. The COMS propulsion system provides the thrust and torque required for the insertion into GEO, attitude and orbit control/adjustment of spacecraft. It is the well-known Chemical Propulsion System(CPS) using bipropellants. On the other hand, the monopropellant propulsion system is employed in KOMPSAT, and its main role is on-station attitude control excluding the orbit transfer function. In this study, these two representative propulsion systems are compared and analysed as well, in terms of essential differences and important characteristics.

• Aerospace Engineering and Technology
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• v.9 no.2
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• pp.14-21
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• 2010

Ion thruster is a kind of electrostatic thruster that use electrostatic field in order to accelerate ionized propellant. Ion thruster have characteristics of small thrust but very high specific impulse among the electric thrusters. High specific impulse can reduce propellant consumption significantly. So, ion thruster have advantage for long time and long distance mission. Recently, plans for space exploration is increasing gradually not only at traditional forward countries for space like USA, Russia and Europe, but also other countries like Japan, China and India. Exploration for superior planets and asteroids the propellant ratio can go up to about 99% when chemical propulsion is used as a cruising thruster. Therefore, latest space exploration vehicles use the ion thruster as main thruster for del-V burn and use monopropellant thrusters for attitude control. In this paper, the development process of preliminary ECR ion thruster and the ECR discharge test results will be presented.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.375-379
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• 2008

In this paper, a 10N class liquid propellant rocket utilizing a dissolving reaction of hydrogen peroxide is constructed and tested. Through a series of designs, seven orifices with a diameter of 200 ${\mu}m$ and a nozzle with a neck of 2.5mm in diameter and area ratio of 2.56 were made. The platinum coated on Isolite was used for catalyst. 90wt% peroxide pressed at 20 bar by nitrogen gas was used for performance evaluation. The length of the catalyst bed and the load of platinum was taken as the parameters for this experiment. For the catalyst support length of 4cm loaded on 5wt% platinum, satisfactory $c^*$ efficiency and stable thrust was observed. The light weight body of the rocket was composed of aluminum. Rocket rose about 10m with relatively constant velocity in launching test.