• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.6
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• pp.19-25
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• 2015

Numerical simulations of the hot gas valve with a pintle have been conducted in order to investigate the effect of numerical methods and computational domains. The grid sensitivity is checked by varying the grid number from 100,000 to 1,700,000. The existence of ambient region doesn't make the significant differences of the flow-field and the temperature distribution. Three turbulence models are adopted to figure out its influence on the thrust and temperature distribution: Spallart-Allmaras, RNG $k-{\varepsilon}$, $k-{\omega}$ SST. The thrusts of the hot gas valve are almost same in all cases of the simulation, however, there are about 5% difference in the temperature distribution. With the ambient region, the difference are observed in the temperature distribution with respect to the number of grids.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.3
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• pp.77-81
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• 2012

In a cryogenic propellant tank the pressurant is contracted due to heat loss and the propellant itself evaporates. On a restartable propulsion system such phenomena are more intensive because the propellant contacts with the pressurant on the larger surface during the coast flight. Such heat and mass transfer phenomena should be considered for estimating the amount of pressurant. On the hypothesis that the heat and mass transfer quasi-equilibrium is achieved during the coast flight, the calculation process of the equilibrium pressure is presented. On the process the amount of loaded helium on the Falcon-1 second stage is calculated.

• 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.34 no.11
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• pp.54-60
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• 2006

This paper discusses wheel offloading approaches of the COMS which has a single solar array system for the accommodation of the optical payloads. First of all, in an effort to reduce fuel consumption and reflect practical implementation point of view, thruster sets for wheel offloading are proposed based on numerical analyses taking into account the COMS configuration. In this analysis, it is assumed that the wheel offloading is conducted twice a day. Secondly, in order to evaluate the effectiveness of the proposed thruster sets, orbit simulations are conducted for several wheel offloading approaches and compared.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.1
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• pp.1-11
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• 2012

Performance analysis was performed for an autorotating rotor. For a given airspeed, shaft angle, and collective pitch, the steady state of autorotation was judged by using the transient simulation method(TSM), then the thrust, lift, and drag coefficient for that state were computed. Average thrust was calculated from the instantaneous thrusts, in which the TSM was used in blade thrust integration. The analysis method was applied to the model rotor that had been tested by wind tunnel. Some comparison between analysis and test was provided. Two types of two-dimensional airfoil aerodynamic data were utilized in analysis, and they were made by Navier-Stokes Solver in terms of Reynolds and Reynolds-Mach number. The quantitative difference of results using two data set was examined and compared.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.196-200
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• 2005

A theoretical model for the calculation of chemical equilibrium composition of propellant combustion product is briefly presented for the performance analysis of monopropellant hydrazine rocket engine. Analysis result is compared to that of test and evaluation of 1-lbf class thruster and is scrutinized primarily from the view point of ammonia dissociation fraction. Chemical equilibrium composition and average molecular weight is additionally depicted according to the variation of propellant inlet pressures and the varying nozzle area ratio. The theoretical analysis is tried as a way of derivation of design parameters for mid- and large-thrust class of monopropellant rocket engines.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.201-205
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• 2011

In a cryogenic propellant tank the pressurant is contracted due to heat loss and the propellant itself evaporates. On a restartable propulsion system such phenomena are more intensive because the propellant contacts with the pressurant on the larger surface during the coast flight. Such heat and mass transfer phenomena should be considered for estimating the amount of pressurant. On the hypothesis that the heat and mass transfer quasi-equilibrium is achieved during the coast flight, the calculation process of the equilibrium pressure is presented. On the process the amount of loaded helium on the Falcon-1 second stage is calculated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.207-210
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• 2006

Injector is one of the most important elements in Liquid rocket Engine design, and how to arrange these injectors on the head determines the engine performance. In this study, when the swirl injectors are used for the 1st designing of injector head of 170 tonf UDMH-LOX as the propellant of LRE, a distribution relation of the mass flow rate per unit area was calculated from the function of ${\Phi}$, which is related with the mass flow rate characteristics of swirl injector. And the combustion characteristics by circumferential axis were estimated using this relation under the consideration of combustion core and film cooling area.

• Journal of the Korean Institute of Illuminating and Electrical Installation Engineers
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• v.17 no.5
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• pp.141-148
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• 2003

In the recently, the necessity of the hybrid type linear stepping motor(HLSM) of linear motion digital actuator has been increased in the various fields of the automatic control system. The HLSM is directly performed without any converting mechanism. Therefore, the HLSM is better advantaged in the efficiency and economical view than a rotary stepping motor. In this paper, we have designed an optimum tooth shape by the 2D finite element method(FEM) to develop the HLSM with longitudinal flux machine(LFM) type, and calculated the thrust force and normal force. And we have manufactured the prototype of it, and have experimented the thrust force characteristics of it.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.102-105
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• 2007

The performance evaluation of the micro thruster utilizing hydrogen peroxide decomposition is described. The catalyst bed was made of porous ceramic material($Isolite^{(R)}$) with large surface to mass ratio. 14%wt platinum was loaded on the catalyst support as a catalyst. Hydrogen peroxide with 85% concentration was used as a monopropellant. The length of the catalyst bed and the feed pressure of the hydrogen peroxide were taken as the parameters for the experiment. All experiments were carried out under cold start condition for 30 seconds. The $c^*$ efficiency was evaluated for each test case using measured pressure data. For the catalyst support length of 30 mm and feed pressure at 5.51 bar, satisfactory $c^*$ efficiency beyond 95% was observed.