• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.6
• /
• pp.59-67
• /
• 2014

Acceptance level random vibration and life firing test for development model of 1 N-class monopropellant thruster have been performed. From the results of random vibration, the natural frequency of the dual thurst module composed of 1 N-class development model thrusters was higher than the part level requirement(>100 Hz) and the structural robustness was verified. Thrust decrease of steady sate was below 7% and thrust instability was within ${\pm}5%$ in the life firing test using over 20 kg propellant throughput. The computerized tomography for catalyst bed showed a less than 7% of catalyst loss and it revealed the design appropriateness of the current thruster development model.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.47 no.5
• /
• pp.335-343
• /
• 2019

The third stage of the KSLV-II is equipped with the reaction control system that performs three axis-control during non-thrust coasting phase and performs a roll axis control during thrust phase. Toxic propellants such as hydrazine have been used for conventional rocket propulsions, however, recently, more studies have been conducted on the use of non-toxic eco-friendly propellants such as ADN and HAN. Especially, hydrogen peroxide has received a growing focus as an emerging propellant. It is considered an alternative of the toxic propellants because of economic advantage in producing the system, conducting operation test, and evaluation of the test result. In this paper, we describes the design, prototype, testing and evaluation of the test results with the 50 N-level hydrogen peroxide monopropellant thruster system which is currently under development.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.25 no.5
• /
• pp.36-52
• /
• 2021

As the application of nano-satellites constellation increases worldwide in the wake of New Space era, there is growing demand for the development of thrusters for precise attitude and orbit control of small satellites. Field Emission Electric Propulsion(FEEP) thruster uses a liquid metal as a propellant and accelerates the ionized liquid metal through a strong electric field at the tip of the metal surface. FEEP thruster technology is suitable for nano-satellites which require various missions for attitude and orbit control, because it provides thrust ranging from 1 µN to 1 mN with high specific impulse up to about 10,000 s and can be miniaturized due to its simple structure. In this paper, the basics of FEEP thrusters are introduced, then the current status of research and development of FEEP thrusters are presented.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.31 no.4
• /
• pp.69-75
• /
• 2003

The ideal thrust function is used to determine the local thrust of supersonic combustor. Method of thrust determination from measured pressures are applied to the Mach 2.5 model supersonic combustor. In this application, measured pressures from the experiments in the University of Michigan are used to determine the local thrust of supersonic combustor. Marginal results of local thrust are obtained and discussed. Combustion and wedge affect thrust distributions in the upstream region significantly. The thrust determination from pressure measurements can be a simple, feasible and applicable method, especially when thrust stand is not available.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.34 no.1
• /
• pp.95-102
• /
• 2006

The PPT being currently developed for the flight model represents a significant leap in techniques and technology compared to the previous flight ones. The electrical energy to be charged in the pulsed plasma thruster (PPT) is a very important aspect to provide an uniform impulse bit ,, and a specific impulse ,, for satellite attitude control. In this paper, we propose a nonlinear control technique and a stability analysis based on the Lyapunov function for the pulsed plasma thruster. Specifically, the proposed control law guarantees to charge and discharge the electrical energy generated from the power processing unit (PPU) within the specified time.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.8 no.4
• /
• pp.84-90
• /
• 2004

A performance evaluation is made in terms of thrust, impulse bit. and specific impulses for a set of mono-propellant hydrazine thrusters producing 0.95 lbf of nominal thrust at an inlet pressure of 350 psia. With a brief description on the hot-firing test configuration and procedures. a typical data obtained from steady-state firing mode is given directly showing the variational behavior of propellant supply pressure, mass flow rate, vacuum condition, and thrust. The performance features are successfully compared to the reference criteria of 1-lbf standard mono-propellant rocket engine. Additionally. a statistical inter-thruster treatment is concisely depicted for the justification of selected thrusters as a grouped member of flight model for spacecraft propulsion system.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.34 no.10
• /
• pp.92-98
• /
• 2006

An 1 lbf of NASA standard monopropellant thruster, MRE-1, is used for KOMPSAT (Korea Multi-Purpose Satellite) which is launched in 2006 and provides reliable and cost-effective means for attitude and maneuvering control system. The monopropellant thruster obtains required thrust by thermal decomposition process of propellant through catalyst bed. During firing, the decomposition plays a role of a heat source that may occur an excessive radiation heat transfer to peripheral structures and electronics in relatively low temperature condition.Therefore, the radiation heat shield is needed to prevent the critical radiative heat exchange between thruster and satellite during firing. This paper summarizes an overall development process of radiation heat shield from the design engineering up to the manufacturing.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.48 no.8
• /
• pp.611-620
• /
• 2020

The effect of discharge voltage on electron mean energy, electric potential, ionization rate, neutral and ion density of Hall thruster was analyzed using a two-dimensional axisymmetric hybrid model. The results of the code developed for this study such as discharge current, thrust, and plasma distribution according to discharge voltage of SPT-100ML Hall thruster were compared by experiments and calculations of other researchers for validation. The results show that the electron mean energy, the ionization rate, and the ion density are increased while the neutral density is decreased as the discharge voltage is increased. The thrust and the discharge current are proportional to the discharge voltage.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.1
• /
• pp.42-49
• /
• 2014

Performance evaluation was carried out for the 70 N-class hydrazine thruster whose design performance had been already verified. The pulse-mode firing test was conducted for the development model thrusters with various thrust chamber diameters. Evaluation was made by the performance parameters such as specific impulse, impulse bit, and characteristic velocity, etc: specific impulse and characteristic velocity were deteriorated as the thrust chamber diameter deviates from a standard model. Consequently, it is revealed that the performance characteristics of standard model is most superior among the test models.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.48 no.7
• /
• pp.529-536
• /
• 2020

Reaction Control System of the third stage of the Korean Space Launch Vehicle II conducts roll control and 3 axis control throughout third stage engine start, satellite separation, and collision and contamination avoidance maneuver. Reaction control system consumes its propellant in each thruster operation. Hence, loading of proper amount of the propellant is important for mission success. It is needed to have a rough estimation method of propellant consumption during the flight. In this paper, we developed a energy equation using pressure and temperature data which are acquired in the on-board reaction control system. We constructed a test system which is similar with the on-board reaction control system to verify the energy equation. Test results using deionized water were compared with estimated propellant consumption. We also conducted an error analysis of the energy equation. We also presented the propellant consumption result of a system level operation test.