• 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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• 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.

• 한국전산유체공학회:학술대회논문집
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• 2008.03b
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• pp.401-404
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• 2008

Pressurization system in a liquid-propellant launcher supplies the controlled gas into the ullage volume of propellant tanks to feed propellants to combustion chamber by pressurizing propellants stored in propellant tanks. The ullage part of propellant tank should be constantly pressurized to supply the propellants stored in propellant tanks to turbo-pump or combustion chamber by pressurant pressurization system. Pressurant used to pressurize propellants is generally stored in a series of tanks at cryogenic temperature and high preassure inside an oxidizer tank. The reason is to store the quantity of pressurant as much as possible and to make pressurant tanks as small as (i.e. as light as) possible. However for test convenience pressurant tank is located at STP (standard temperature and pressure) environment in this study. Orifices are widely adapted to several pressurization systems in liquid rocket propulsion systems. Discharge coefficients of orifices are essentially needed for the optimized design of pressurization system in liquid rocket propulsion system. For this study gaseous nitrogen was served as pressurant and rounded entry orifices were employed. The forty-two (42) rounded entry orifices (the radii of curvatures are 0.5 and 1.0) have been tested experimentally in the supersonic flow region. The discharge coefficients of rounded entry orifices with inside diameters ranging from about 1.4 to 5.0mm was measured with 0.95 ${\sim}$ 0.99.

• Aerospace Engineering and Technology
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• v.10 no.1
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• pp.54-62
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• 2011

The pressurant mass required for propellant tank pressurization with operating condition variation was estimated by using the numerical model already developed for this purpose. The model was applied to the concept design results of KSLV-II first stage oxygen tank. The supplied pressurant temperature, oxygen volumetric flow rate, and the ratio of length to diameter of the tank were selected as variables. The required pressurant mass and mass flow rate, collapse factor, ullage temperature distribution were predicted, and the results showed that the pressurant temperature had the largest effect on the amount of the required pressurant mass. The pressurizing efficiency of the propellant tank was calculated through analyzing energy distribution in the ullage. It was found that the gas-to-wall heat transfer in the ullage was dominant, and much of the pressurant energy was lost to tank wall heating.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.1-10
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• 2006

In this paper, a calculation method of the thermodynamic parameters of cryogenic propellant is proposed when a cryogenic propellant tank is pressurized by gaseous helium(GHe) bubbling. Temperature of cryogenic propellant and mass of dissolved GHe into propellant were analyzed at the various operation of pressurization of tile liquid oxygen(LOX) and hydrogen($LH_2$) tank using helium bubbling. It was evaluated how the GHe bubbling influences to the thermodynamic parameters of LOX and $LH_2$ with results of the analysis. With the proposed calculation method, It will be able to confirm the feasibility of GHe bubbling as a pressurization system of cryogenic propellant tank and to optimize the pressurization system using GHe bubbling.

• Progress in Superconductivity and Cryogenics
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• v.16 no.4
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• pp.71-77
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• 2014

For a long-term space mission, filling process of cryogenic liquid propellant is operated on a space vehicle in space. A vent process during transfer and filling of cryogenic propellant is needed to maintain the fuel tank pressure at a safe level due to its volatile characteristic. It is possible that both liquid and vapor phases of the cryogenic propellant are released simultaneously to outer space when the vent process occurs under low gravity environment. As a result, the existing filling process with venting not only accompanies wasting liquid propellant, but also consumes extra fuel to compensate for the unexpected momentum originated from the vent process. No-Vent Fill (NVF) method, a filling procedure without a venting process of cryogenic liquid propellant, is an attractive technology to perform a long-term space mission. In this paper, the preliminary experimental results of the NVF process are described. The experimental set-up consists of a 9-liter cryogenic liquid receiver tank and a supply tank. Liquid nitrogen ($LN_2$) is used to simulate the behavior of cryogenic propellant. The whole situation in the receiver tank during NVF is monitored. The major experimental parameter in the experiment is the mass flow rate of the liquid nitrogen. The experimental results demonstrate that as the mass flow rate is increased, NVF process is conducted successfully. The quality and the inlet temperature of the injected $LN_2$ are affected by the mass flow rate. These parameters determine success of NVF.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.4
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• pp.50-58
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• 2016

Korean Space Launch Vehicle (KSLV-II) Propellant Supply System charges liquid oxygen and kerosene to each propellant tank for the stages. To charge the launch vehicle propellant tank safety, the propellant charge flow rates and scenarios should be defined. First, the Propellant Supply System was modeled with 1D flow analysis program. The control valve capacity and orifice size were calculated by performing the 1D steady state simulation. Second, the 1D transient simulation was performed by using the steady state simulation results. As propellants were being charged at the each tank, the increased tank liquid level decreases the charge flow rate. Consequently, the proposed supply system satisfies the required design charging conditions.

• Progress in Superconductivity and Cryogenics
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• v.16 no.2
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• pp.64-69
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• 2014

Measuring an exact amount of remaining cryogenic liquid propellant under microgravity condition is one of the important issues of rocket vehicle. A Pressure-Volume-Temperature (PVT) gauging method is attractive due to its minimal additional hardware and simple gauging process. In this paper, PVT gauging method using liquid nitrogen is investigated under microgravity condition with parabolic flight. A 9.2 litre metal cryogenic liquid storage tank containing approximately 30% of liquid nitrogen is pressurized by ambient temperature helium gas. During microgravity condition, the inside of the liquid tank becomes near-isothermal condition within 1 K difference indicated by 6 silicon diode sensors vertically distributed in the middle of the liquid tank. Helium injection with higher mass flow rate after 10 seconds of the waiting time results in successful measurements of helium partial pressure in the tank. Average liquid volume measurement error is within 11% of the whole liquid tank volume and standard deviation of errors is 11.9. As a result, the applicability of PVT gauging method to liquid propellant stored in space is proven with good measurement accuracy.

• 한국전산유체공학회:학술대회논문집
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• 2004.10a
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• pp.131-136
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• 2004

Thermal control of satellite propellant tank is achieved by patch heaters enabled by thermostat's behavior. It is important to attach the thermostat on the appropriate position of the propellant tank. However its position cannot be given with exact numerics because tank is spherical. Actually the position for thermostat is designated in relevant drawing approximately, therby, the engineer practices depending on his own experience and intuition. The sensitivity analysis for the position of thermostat is performed such that the influence on the thermal behavior and control of tank is examined quantatively. When assembling tank module, the reasonable performance on the thermal control is believed with possible human errors if the uncertainty in the position of thermostat is not quite large.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.813-816
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• 2011

The actuation and leakage tests of solenoid valve for propellant tank pressurization system have been conducted. The response time of solenoid valve manufactured is satisfactory to perform requirement. However, leakage was found at the upper part seat of relief valve inside solenoid valve. Solenoid valve was disassembled in order to discover leakage causes. We found out that the upper seat of relief valve was damaged. Through this study, the development possibility of propellant tank pressurization solenoid valve was confirmed.