• 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.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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• 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.

• 한국전산유체공학회:학술대회논문집
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• 2005.04a
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• pp.96-99
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• 2005

The geostationary satellite propulsion system has thermistors which can measure liquid propellant temperature at tanks, pipes and etc. In the satellite propulsion system with several tanks, the propellant in the tanks is moved by temperature change and this temperature pattern is constant. In this paper, the temperature change pattern of KOREASAT 1 propulsion system is compared and the prediction study of pressurant inflow using temperature change of geostationary satellite propulsion system is described.

• Proceedings of the KSRS Conference
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• 2007.10a
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• pp.179-182
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• 2007

COMS (Communication, Ocean and Meteorological Satellite) is a geosynchronous satellite and has been developing by KARI and Astrium for Ka-band communication, ocean observation and meteorological observation. COMS Chemical Propulsion System (CPS) uses a bipropellant propulsion subsystem, which is applied for transferring COMS from GTO to GEO (mission orbit) and implementing station-keeping manoeuvres. In this paper COMS CPS is briefly introduced for understanding. A few of mathematical thermodynamic modelings of bipropellant propulsion system in literatures are reviewed and authors has studied those models for developing a computer program, which predicts variations of thermodynamic properties such as temperature and pressure histories in the helium pressurant tank, MMH propellant tank and NTO propellant tank during LAE firing and on-orbit manoeuvrings. The CPS thermodynamic model may be used to compute pressurant and propellant masses and to size tank volumes.

• Aerospace Engineering and Technology
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• v.1 no.1
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• pp.55-64
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• 2002

A conceptual design of propulsion system for a geosynchronous communication satellite with 12 years design life is presented in this paper. Propellant mass budget for the design life is calculated using total velocity increment (ΔV) flowed-down from mission requirement analysis. Sizes of the fuel and oxidizer tank are derived based on the calculated propellant mass budget, and mass of the pressurant as well as the size and pressure of pressurant tank are calculated too. Thruster positioning, number of rocket engines, and position of tank are determined through Trade-Off Study with Structure & Mechanical Subsystem. Propulsion system configuration and its schematics are presented finally.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.721-724
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• 2011

A system to pressurize propellants stored in propellant tanks is necessary to feed liquid-propellants into combustion devices at the required pressure and flowrate without having cavitation in turbo-pumps. A pressurization system can be categorized into pre-pressurization stage and main-pressurization stage. This report is regarding to a main-pressurization system. Pressurization methods for propellant tanks are divided into pressurant gas generating method and pressurant gas feeding method. One of pressurant gas generating methods uses the vaporized oxygen gas from cryogenic liquid oxygen and non-flammable gas. In this report, both advantages and disadvantages for pressurization methods and types of pressurization systems are compared. Especially the characteristics and principle of pressurization system using impulsive control strategy applied in launch vehicles are introduced. Additionally the structure, schematics, and specifications of heat exchanger, which is one of main units in pressurization system are also discussed. This paper can be utilized to generate the conceptual requirements and to design preliminary configuration of pressurization system during the development of launch vehicle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.168-175
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• 2005

The pressurization system in a liquid rocket propulsion system provides a controlled gas pressure in the ullage space of the vehicle propellant tanks. It is advantage to employ a hot gas heat exchanger in the pressurization system to increase the specific volume of the pressurant and thereby reduce over-all system weight. A significant improvement in pressurization-system performance can be achieved, particularly in a cryogenic system, where the gas supply is stored inside the cryogenic propellant tank. The temperature characteristic of cryogenic pressurant is very important to develop some components in pressurization system. Numerical modeling and Test data were studied using SINDA/FLUINT Program and PTF(Propellant-feeding Test Facility).