• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.984-990
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• 2017

Mesh screen modeling and liquid propellant discharge simulation of surface tension tank were performed using commercial CFD software Flow-3d. $350{\times}2600$, $400{\times}3000$ and $510{\times}3600$ DTW mesh screen were modeled using macroscopic porous media model. Porosity, capillary pressure, and drag coefficient were assigned for each mesh screen model, and bubble point simulations were performed. The mesh screen model was validated with the experimental data. Based on the screen modeling, liquid propellant discharge simulation from PMD tank was performed. NTO was assigned as the liquid propellant, and void was set to flow into the tank inlet to achieve an initial volume flow rate of liquid propellant in $3{\times}10^{-3}g$ acceleration condition. The intial flow pressure drop through the mesh screen was approximately 270 Pa, and the pressure drop increased with time. Liquid propellant discharge was sustained until the flow pressure drop reached approximately 630 Pa, which was near the estimated bubble point value of the screen model.

• Aerospace Engineering and Technology
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• v.12 no.2
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• pp.83-90
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• 2013

In this paper, the energy flows related to cryogenic propellant tank ullage were understood and pressurization efficiency of the tank was calculated using propellant feeding test data with the help of calculation program. The related energy flow terms and calculation method of each terms were described. Three test data of different tank pressure and incoming pressurant temperature were used. Under the test conditions, the pressurization efficiency was low in the range of 13.9%~19.3%. The proportion of energy loss to the incoming pressurant energy was in the range of 55.2%~67.6%. The energy loss to the propellant tank wall was the biggest one. If the temperature of incoming pressurant was the same, the rates of each energy flows to the incoming energy were almost the same regardless of the propellant tank pressure. The collapse factor of propellant tank was calculated using test data, and the relation of it to the heat loss rate was observed.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.1
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• pp.7-15
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• 2018

In this study, AOT that is used as a surfactant in various industries was applied to an HTPB/AP solid propellant. AOT is one of the anionic surfactants, and there have been cases where AOT was reported to induce self-extinguishable properties in propellants overseas. In this study, solid propellants using AOT were prepared, and their properties and combustion characteristics were investigated. The combustion rate of the AOT-applied propellant drops sharply when the pressure reaches a certain value during combustion. Further, the density and hardness of the propellant are lower than those of conventional HTPB/AP propellants.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.2
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• pp.181-188
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• 2016

It is known that reliability of liquid rocket engines depends on the design thrust, propellant, engine cycle, and hot firing test time. Previously, a method was developed for estimating reliability of a new engine by adjusting the design thrust and hot firing test time of reference engines where reference engines have the same propellant and engine cycle with the new engine. In this paper, we provide a procedure to predict the engine reliability when the new engine and the reference engine have different propellant and engine cycles. The proposed method is illustrated to estimate the engine reliability of the first stage of Korea Space Launch Vehicle II.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.793-794
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• 2011

In launch process, propellants should be supplied with established temperature range for engine normal operation. In order to satisfy this temperature condition, propellant feeding systems should be considered some effects during operation. This paper studied liquid oxygen filling system operation process and cooling method of liquid oxygen during launch process.

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.85-89
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• 2006

Propulsion System provides the required velocity change impulse for orbit transfer from parking orbit to mission orbit and three-axis vehicle attitude control impulse. New LEO Satellite propulsion system (PS) will be an all-welded, monopropellant hydrazine system. The PS consists of the subassemblies and components such as Thrusters, Propellant Tank, Pressure Transducer, Propellant Filter, Latching Isolation Valves, Fill/Drain Valves, interconnecting propellant line assembly, and thermal hardwares for operation-environment control of the PS. In this study, preliminary design process of LEO Satellite propulsion system will be summarized.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.141-145
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• 2004

The residual propellant of satellite is the primary factor of satellite life. This propellant used in the satellite is stored as liquid in tanks. But it is very difficult to accurately measure propellant to be used for maintaining of satellite by an irregular influence of environment. In this paper, a new method of gauging propellant residual of satellite employing multi-tank system by measuring mass flow of thermal pumping liquid propellant is presented. In cases of being connected between tanks, propellant in tanks move by temperature difference of tanks. If propellant mass flow is measured at line between tanks, residual propellant in tanks is able to be estimated.

• 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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• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.4
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• pp.52-62
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• 2017

The study for the performance characteristics of the nozzleless booster used in ramjet booster was carried out. Performances related to pressure and thrust for nozzleless booster are lower than classical motor those because of absence of convergent and divergent sections of nozzle. To solve this problem, it developed a high-performance propellant with maximum impulse density included Al as metal fuel. Using the nozzleless booster casted the propellant, ground test of it was carried out by varying the length-to-diameter ratio (L/D ratio) of the propellant. Specific impulse of nozzleless booster was limited to about 75 percents of its value compared with that of classical motor adapted nozzle in the same propellant and propellant length and will be estimated approximately 85 percents of its value compared with that of classical motor at same average pressure in terms of the curve fitting by our test results.