• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.3
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• pp.54-62
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• 2016

Micro plasma thruster (${\mu}PT$) was studied experimentally with a dual-stage micro-hollow cathode discharge (MHCD) plasma. Electrostatic-like acceleration exhibiting more directional and elongated exhaust plume was achieved by a dual layer MHCD at the total input power less than 10 W with argon flow rate of 40 sccm. V-I characteristic indicated that there was an optimal regime for dual-stage operation where the acceleration voltage across the second stage remained constant. Estimated exhaust plume length showed a similar trend to the analytic estimate of exhaust velocity which scales with an acceleration voltage. ${\mu}PT$ with multiple holes exhibited similar performance with single-hole thruster indicating that higher power loading is possible owing to decreased power through each hole. Boltzmann plot of atomic argon spectral lines showed average electron excitation temperature of about 2.6 eV (~30,170 K) in the exhaust plume.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.6
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• pp.126-133
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• 2004

MRE-1 Dual Thruster Module(DTM), which will be used in KOMPSAT(Korea Multi-Purpose Satellite), can provide reliable and cost-effective means for attitude and maneuvering control system. Thruster heat shield, one of the main components of DTM, is designed to prevent the critical radiative heat exchange between thruster and satellite during firing. To overcome the manufacturing difficulties, a electroforming process is preferred to classical welding process. In this case, an inner diameter of a new shield will be decreased a little due to the change of manufacturing process. Therefore, the interference problem between thruster nozzle and heat shield is investigated through structural analysis and their results are described in this paper.

• 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.

• Aerospace Engineering and Technology
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• v.8 no.1
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• pp.144-153
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• 2009

To study the plume effects in the rarefied region, the Direct Simulation Monte Carlo(DSMC) method is usually adopted because the plume field usually contains the entire range of flow regime from the near-continuum in the vicinity of nozzle exit through transitional state to free molecular at far field region from the nozzle. The objective of this study is to investigate the behaviors of a small monopropellant thruster plume in the rarefied region numerically using DSMC method. To deduce accurate results efficiently, the preconditioned scheme is introduced to calculate continuum flow fields inside thruster to predict nozzle exit properties used for inlet conditions of DSMC method. By combining these two methods, the rarefied flow characteristics of plume such as strong nonequilibrium near nozzle exit, large back flow region, etc, can be investigated.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.9
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• pp.906-915
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• 2009

To study the plume effects in the vacuum area, the Direct Simulation Monte Carlo(DSMC) method is usually adopted because the plume field usually contains the entire range of flow regime from the near-continuum in the vicinity of nozzle exit through transitional state to free molecular at far field region from the nozzle. The objective of this study is to investigate the behaviors of a small monopropellant thruster plume in the vacuum area numerically using DSMC method. To deduce accurate results efficiently, the preconditioned scheme is introduced to calculate continuum flow fields inside thruster to predict nozzle exit properties used for inlet conditions of DSMC method. By combining these two methods, the vacuum flow characteristics of plume such as strong nonequilibrium near nozzle exit, large back flow area, etc, can be investigated.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.9
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• pp.799-804
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• 2015

The purpose of this study is to analyse the dynamic disturbances(disturbed forces and disturbed torques) due to attitude control thruster's plume impingement on the solar arrays. To produce database of the dynamic disturbances a sweep analysis was done, in which the two parameters are used; the distance between the thruster and solar arrays and the thruster tilt angle. Based on the database, a third order polynomial approximation is computed to represent the characteristics of the disturbed forces and torques. The final results are the coefficients of the approximation for each solar array angle position. These results as input data are used to optimize the configuration of the attitude control thrusters. This analysis is appled to the two candidate solar arrays for Geo-Kompsat-2 satellite and the results of the disturbed forces and disturbed torques are compared and analysed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.66-69
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• 2006

A two-dimensional Particle-In-Cell (PIC) simulation of a Hall thruster is presented. The thruster is being developed for orbit transfer and correction of a small satellite. Preliminary investigation of the simulation result finds well separated acceleration and ionization layers. The simulation further shows that collisional ionization of the xenon neutrals allows sufficient acceleration of the ionized plasmas that is adequate for the intended correction and transfer of small satellite orbits. Anticipated performance of the thruster based upon the present results will be calculated.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.4
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• pp.300-309
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• 2017

A spacecraft obtains a reaction momentum required for an orbit correction and an attitude control by exhausting a combustion gas through a small thruster in space. If the exhaust plume collides with spacecraft surfaces, it is very important to predict the exhaust plume behavior of the thruster when designing a satellite, because a generated disturbance force/torque, a heat load and a surface contamination can yield a life shortening and a reduction of the spacecraft function. The purpose of the present study is to ensure the core technology required for the spacecraft design by analyzing numerically the exhaust gas behavior of the 10 N class bipropellant thruster for an attitude control of the spacecraft. To do this, calculation results of chemical equilibrium reaction between a MMH for fuel and a NTO for oxidizer, and continuum region of the nozzle inside are implemented as inlet conditions of the DSMC method for the exhaust plume analysis. From these results, it is possible to predict a nonequilibrium expansion such as a species separation and a backflow in the vicinity of the bipropellant thruster nozzle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.929-932
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• 2011

We developed a valve actuator controller for thruster system. This thruster system has four actuators and the actuator use a BLDC motor. Controller was made based on system and control requirement. The controller is consist of power, control and Amp. The control module use a micro-controller which is TMS320F28335 of Texas Instruments. It works for digital PID control and CAN communication and system control. The amp module for three phase BLDC motor use IGBT.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.35-38
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• 2005

The computational fluid dynamic analysis has been conducted for the thermo-chemical flow field in an arcjet thruster with mono-propellant hydrazine ($N_{2}H_4$) as a working fluid. Coupled Reynolds Averaged Navier-Stokes (RANS) equations and Maxwell equations were used to account for the Ohm heating and Lorentz forces. Hydrazine chemistry and thermal radiation were also incorporated to the fluid dynamic equations by assuming infinitely-fast reactions and optically thick media. In addition to the thermo-physical understandings of the flow field inside the arcjet thruster, results shows that performance indices are improved by amount of $20\%$ in thrust and $200\%$ in specific impulse with the 0.6kW are heating.