• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.2
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• pp.63-70
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• 2013

To analyze flow characteristics of the side jet thruster with 4 shutters and rectangular nozzles, a 3-D simulation has been implemented. Numerical calculations for two rotation angles of the shutter, have been conducted. Internal recirculation in a chamber and asymmetric flow structure in a nozzle were observed. In addition, the more shutter rotated, the more asymmetries of flow increased, and this phenomena resulted in thrust bias. The degrees of thrust bias and thrust performance with the rotation angles of the shutter were predicted and compared with theoretical thrust.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.4
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• pp.1-11
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• 2020

In this study, study on the internal ballistic analysis method for single-chamber dual-thrust rocket motors meeting a dual-thrust profile requirement by tailoring the grain burning area is presented. The analysis method, which can acquire variables required for the performance prediction, considering gradual change of burning rate correction factor and specific impulse in the transition phase, is proposed. Improvements compared to the analysis method in the previous study, which do not consider change in the transition phase, are verified through comparison between the newly proposed method and the method in the previous study. Internal ballistic variables are obtained for four different ground firing test conditions using the proposed method, and the performance prediction for each condition is conducted using these variables. These prediction results and the ground test data are in good agreement, so it is confirmed that the performance prediction of dual-thrust motors with same design geometries based on the proposed analysis method is available.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.5
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• pp.29-36
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• 2012

3-D linear acoustic analysis has been performed to elucidate damping characteristics of large Helmholtz acoustic cavities in FASTRAC thrust chamber. Acoustic impedance concept has been introduced to quantify combustion stabilization capacity. For a given acoustic cavity, sonic velocity in the cavity to achieve an optimal tuning has been determined and satisfactory agreement with the previous results has been obtained. Feasible estimation of sonic velocity in the acoustic cavity has been devised. Results show similar trends without significant deviations, which can be used in the procedure of design and verification of acoustic cavity. From the satisfactory results, investigation of other thrust chambers with acoustic cavities which have shown combustion instabilities will be done as future works.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.2
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• pp.25-33
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• 2012

A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Pressure of the main combustion chamber, nozzle expansion ratio and O/F ratio have been selected as design variables. The target engine has the open gas generator cycle using the LO2/RP-1 propellant. The gas properties of the combustion chamber have been obtained from CEA2 and the mass has been estimated using reference data. The objective function has been set as multi-objective function with the specific impulse and thrust to weight ratio using the weight method. The result shows about 4% improvement of the specific impulse and 23% increase of the thrust to weight ratio. The Pareto frontier line has been also obtained for various thrust requirements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.123-127
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• 2005

LRE(liquid rocket engine) performance design code with several modules for each engine component has been developed for a preliminary design purpose. Thrust chamber, non-cryogenic centrifugal pump, single stage axial impulse turbine, gas generator and exhaust pipe for extra thrust have been considered. For simplicity, pump exit pressures are fixed, which eliminates pressure balancing problem between thrust chamber and turbopump unit. In this paper, calculated performance parameters with system flow charts and the design methodologies for each component are briefly presented and the results are compared with tile real engine specification.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.10
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• pp.985-991
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• 2010

The objective of this work was to investigate the pintle-tip shape effect on nozzle flow and thrust by cold flow test. When nozzle throat area was decreased by pintle movement, chamber pressure was increased monotonously but thrust was increased differently according to every pintle-tip shape. At the same chamber pressure and nozzle throat area, thrust of convex pintle-tip shape was mostly larger than that of concave one. Nozzle wall pressure distribution and magnitude of pintle-tip load depended on the pintle-tip shape, pintle position and nozzle throat area.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.942-948
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• 2011

Solid propulsion systems have simple structures compared to other propulsion systems and are suitable for long-term storage. However the systems generally have limits on control of thrust levels. In this paper we introduce controllers for combustion chamber pressure using on-off control techniques which have been known for relatively easy implementation and energy efficiency. For this, we use a simple pressure change model by considering only mass conservation within the combustion chamber and we design a classical controllers and on-off controllers with are Pulse Width Modulation(PWM) and Pulse Width Pulse Frequency Modulation (PWPFM). Then we compare the performance results of the controllers through numerical simulations.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.127-134
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• 2011

A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Pressure of the main combustion chamber, nozzle expansion ratio and O/F ratio have been selected as design variables. The target engine has the open gas generator cycle using the LO2/RP-1 propellant. The gas properties of the combustion chamber have been obtained from CEA2 and the mass has been estimated using reference data. The objective function has been set as multi-objective function with the specific impulse and thrust to weight ratio using the weight method. The result shows about 4% improvement of the specific impulse and 23% increase of the thrust to weight ratio. The Pareto frontier line has been also obtained for various thrust requirements.

• Aerospace Engineering and Technology
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• v.1 no.2
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• pp.51-56
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• 2002

Satellite propulsion system is employed for orbit transfer, orbit correction, and attitude control. The monopropellant feeding system in the low-earth-orbit satellite blowdowns fuel to the thrust chamber. The thrust produced by the thruster depends on fuel amount flowed into the combustion chamber. If the thruster valve be given on-off signal from on-board commander in the satellite, valve will be opened or closed. When the thrusters fire fuel flows through opened thruster valve, instantaneous stoppage of flow in according to valve actuation produces transient pressure due to pressure wave. This paper describes transient pressure predictions of the KOMPSAT-2 propulsion system resulting from latching valve and thrust control valve operations. The time-dependent set of the fluid mass and momentum equations are calculated by MOC.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.125-130
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• 2004

Experimental study on combustion characteristics of double swirl coaxial injectors has been conducted for the assessment of critical injector design parameters. A reusable, unielement thrust chamber has been fabricated with a water-cooled copper nozzle. Two principle design parameters, a swirl angle and a recess length, have been investigated through hot firing tests for the understanding of their effects on high pressure combustion. Clearly, both parameters considerably affect the combustion efficiency, dynamics and hydraulic characteristics of an injector. Internal mixing of propellants in a recess region increases combustion efficiency along with the increase of a pressure drop required for flowing the same amount of mass flow rates. It is concluded that pressure buildup due to flame can be released by the increase of LOx flow axial momentum or the reduction of a recess length. Dynamic pressure measurements of the thrust chamber show varied dynamic behaviors depending on injector configurations.