• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.723-725
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• 2010

Calculation program to predict required pressurant mass for propellant tank was verified through flight test data. This program was already developed and verified through ground test data, but to increase reliability of program, it was compared with flight test data of KSR-III launched in 2002. Because pressurant temperature incoming to propellant tank was not measured in flight test, that was assumed in calculation program. Required pressurant mass and inside temperature of oxygen tank dome was compared. Validation of calculation program was verified by showing required pressurant mass accuracy of 6%.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.293-297
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• 2009

Valves, which are used to supply or block the flow of cryogenic pressurant in the pressurization system of liquid-propellant propulsion system in a launch vehicle, are pneumo-actuated valve, solenoid valve, pyro-valve, etc. Both pneumo-actuated valve and solenoid valve have more complex structure and are heavier than pyro-valve. For this study, a couple of pyro-valves, which are applicable to cryogenic and high-pressure fluid (cryogenic gaseous helium), have a simple structure, and are comparably light, are designed, manufactured, and tested (proof-pressure/leakage tests, performance test, vibration test, helium supply tests).

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.1
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• pp.73-79
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• 2010

Propellant tank pressurizing gas injector is used in the pressurization system of liquid propellant rocket to reduce incoming gas velocity and distribute the gas in the tank. Temperature distribution in the propellant tank ullage is varied according to the gas injector shape, and it has influence on the required pressurant gas and thermal phenomena in the tank. In this paper, diffuser type gas injector was studied to make the ullage have stratified temperature distribution. Injected gas flow at the outlet of prototype diffuser was visulized using particle image velocimetry method and it was compared with the results of calculation. Calculation was well agreed with measurement and was used as an inlet condition of propellant tank ullage calculation.

• Journal of computational fluids engineering
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• v.9 no.3
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• pp.49-56
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• 2004

A heated and expanded helium is used to pressurize liquid propellants in propellant tanks of propulsion system of liquid propellant launch vehicles. To produce a heated and expanded helium, an hot-gas heat exchanger is used by utilizing heat source from an exhausted gas, which was generated in a gas generator to operate turbine of turbo-pump and dumped out through an exhaust duct of engine. Both experimental and numerical approaches of hot-gas heat exchanger design were conducted in the present study. Experimentally, siliconites - electrical resistance types - were used to simulate the full heat condition instead of an exhausted gas. Cryogenic heat exchangers, which were immersed in a liquid nitrogen pool, were used to feed cryogenic gaseous helium in a hot-gas heat exchanger. Numerical simulation was made using commercially utilized solver - Fluent V.6.0 - to validate experimental results. Helically coiled stainless steel pipe and stainless steel exhausted duct were consisted of tetrahedron unstructured mesh. Helium was a working fluid Inside helical heat coil and regarded as an ideal gas. Realizable k-』 turbulent modeling was adopted to take turbulent mixing effects in consideration. Comparisons between experimental results and numerical solutions are Presented. It is observed that a resulted hot-gas heat exchanger design is reliable based on the comparison of both results.

• Journal of the Korean Society of Propulsion Engineers
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• v.2 no.1
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• pp.21-30
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• 1998

Design parameters which used to analyze the stress distribution on the tank wall were defined to develop the propellant tank and obtain optimal values. 1/4 modeling of total tank was selected to calculate the stress distribution with respect to the variation of the support lug location and the tank wall thickness and 1/2 modeling was selected for the stress distributions with respect to the variation of fuel outlet location. Actually, 350psi was applied as static load and 12 gravity as a dynamic load during launching on the internal tank wall. The structural analysis was done with respect to attaching condition of the tank in the satellite. Also the effect of the variation of the propellant outlet location from $0^{\cire}$ to $25^{\cire}$ on the stress distribution was investigated. The equivalent stress distribution and optimal parameters induced from analysis results of the each condition will be used as the fundamental data to design the propellant tank.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.363-367
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• 2009

For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.1
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• pp.63-68
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• 2011

For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.3
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• pp.258-263
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• 2012

In this paper the Chemical Propulsion Subsystem of COMS is briefly explained and some telemetries acquired by a series operations of CPS during the Launch and Early Operation Phase of COMS are presented. The pressure and temperature of pressurant tank telemetries are compared with the results of the developed computer program. The changes in pressure are due to the two major phases. The first one is the initialization phases of CPS composed of the venting phase to vent the helium gas in the pipe network from the downstream of the propellant tanks to the thrusters for safety, the priming phase to fill the vented pipe network with oxidizer and fuel respectively and then the pressurization phase to pressurize the ullage of propellant tank to regulated pressure. And the other is the apogee engine firings in which COMS CPS is in the orbit raising phase to use helium as a pressurant to keep the pressure of propellant tank as the liquid apogee engine get fired until COMS reached to the target orbit. This program can be applicable to prepare basis design data of the next Geostationary Satellite CPS.