• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.574-578
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• 2008

In order to achieve thrust vector control and discharge stabilization in Hall thrusters, the azimuthal nonuniformity of propellant flow rate in an acceleration channel was created. A plenum chamber was divided into two rooms by two walls and propellant flow rate supplied to each section was independently controlled. In a magnetic layer type Hall thruster, steering angle of up to ${\pm}2.3$ degree was achieved. In an anode layer type Hall thruster, discharge current oscillation amplitude was decreased with the normalized differential mass flow rate.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.125-129
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• 2009

The basic design of liquid rocket engine combustion chamber for a large space launch vehicle was described. It has vacuum thrust of 74.8 ton, vacuum specific impulse of 306.9 sec, chamber pressure of 60 bar, mass flow rate of 243.6 kg/s and combustion characteristic velocity of 1730 m/sec. The details of combustion performance and geometrical parameter were also given. The 75 ton combustion chamber consists of the combustor head with injector and the chamber/nozzle with regenerative cooling channels.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.10a
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• pp.19-20
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• 2002

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.1
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• pp.79-85
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• 2008

Generally Solid Rocket Motor(SRM) has advantages like this - safety, simplicity and flexibility in design and manufacturing process. However, once propellant grain shape and nozzle throat area are determined, modification of thrust magnitude is nearly impossible. Recently, methods for controlling the thrust magnitude of SRM are vigorously developed. This paper predicts internal ballistic performance variation, especially thrust of SRM by means of Linear Approximation according as chamber pressure or nozzle throat area is changed. The results predicted by the proposed method are good agreement with the those of exclusive Ballistic Performance Prediction Program(SPP).

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.1
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• pp.76-81
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• 2009

A nonlinear pressure controller to actively regulate the thrust of a solid propulsion system is presented. To compensate for the parametric uncertainties with respect to the chamber pressure induced by changing nozzle throat area, Lyapunov-based parameter adaptation method has been applied. In order to verify the effectiveness of the proposed control method, the experiments were carried out using the cold gas test equipment that can simulate the operating environment of variable thrust solid propulsion system. The experiment results show that the nonlinear pressure controller has better performance than conventional P and PI controller.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.370-373
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• 2007

In this study, a hybrid rocket motor with separable and detachable oxidizer tank from combustion chamber is developed. Initially, the measured thrust of the motor showed about 30% of the design thrust since the oxidizer supply was not enough. In order to solve this problem, application is made to expand the orifice diameter of oxidizer injector empirically, so that the mass flow rate of oxidizer was improved. The improved performance was about 60% of design thrust, 18kgf, and thrust-to-weight ratio was reasonable, compared with other sounding rockets.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.48-55
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• 2012

An ambient hot-firing test was carried out for the hydrazine thruster which may be employed in the space launch vehicles. The thruster is designed to produce 67 N (15 $lb_f$) of nominal steady-state thrust at an inlet pressure of 2.41 MPa (350 psia). A scrutiny into the performance characteristics of thruster is made in terms of thrust, propellant supply pressure, mass flow rate, chamber pressure, and temperature at the steady-state firing mode. As a result, it is ensured that the practical performance efficiencies are above 89.1% compared to its ideal requirements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.611-616
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• 2004

In this study, microfabrication of a micro-arcjet nozzle with Fifth-harmonic generation Nd：YAG pulses (wavelength 213 nm) and its thrust performance tests were conducted. A micro-arcjet nozzle was machined in a 1.2 mm thick quartz plate. Sizes of the nozzle were 0.44 mm in width of the nozzle exit and constrictor diameter of 0.1 mm. For an anode, a thin film of Au (~100 nm thick) was deposited by DC discharge PVD in vacuum on divergent part of the nozzle. As for a cathode, an Au film was also coated on inner wall surface. In operational tests, a stable discharge was observed for mass flow of 1.0mg/sec, discharge current of 6 ㎃, discharge voltage of 600 V, or 3.6 W input power (specific power of 3.6 MW/kg). In this case, plenum pressure of the discharge chamber was 80 ㎪. With 3.6 W input power, thrust obtained was 1.4 mN giving specific impulse of 138 sec with thrust efficiency of 24 %.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.78-82
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• 2003

The cooling mechanism for a regenerative cooling liquid rocket engine of 10tf-thrust using kerosene as a fuel was studied from the viewpoint of curtain cooling. Based on the concept of a highly-stratified gas flow in the combustion chamber, the cross section of the combustion chamber was spilt into 2 independent parts, core and exterior part. Additional fuel is injected into the exterior section and gas temperature can be reduced in the exterior section. Consequently, the heat flux into the coolant and wall temperature are reduced and the thermal stability of a liquid rocket engine could be improved.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.10
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• pp.66-72
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• 2003

The cooling mechanism for a liquid rocket engine of 10tf-thrust using kerosene as a fuel was studied from the viewpoint of both the regenerative and curtain cooling. Based on the concept of a highly-stratified gas flow in the combustion chamber, the cross section of the combustion chamber was spilt into 2 independent parts, core and exterior part. Additional fuel is injected into the exterior section and gas temperature can be reduced in the exterior section. Consequently, the heat flux into the coolant and wall temperature are reduced and the thermal stability of a liquid rocket en g i.ne could be improved.