• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.2
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• pp.21-26
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• 2019

Exposure to a rapidly depressurized environment causes extinction of a burning solid propellant. Experiments have been conducted to determine the rate of depressurization required to extinguish a burning solid propellant. For this purpose, a depressurization combustor was designed and fabricated. The results of this experiment were used to determine the boundary between extinction and non-extinction of AP/HTPB solid propellants under different propellant compositions. Experimental results show that the initial and final pressures have a considerable effect on the critical depressurization rate.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.58-64
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• 2007

Propellant pressurization system in liquid rocket propulsion system plays a role supplying pressurant gas at a controlled pressure into the ullage space of propellant tanks. The most important design parameter for such propellant pressurization system is the temperature of pressurant gas fed from pressurant tank, which is placed inside of cryogenic propellant tank. Such pressurant is gaseous state, of which density is very sensitive to the temperature of pressurant. Previous investigation dealt with thermal correlation of pressurant and external fluid at room temperature. This study investigates the temperature variation of cryogenic pressurant (GHe) at the time when the pressurant is coming out of pressurant tank, which is submerged in a liquid oxygen, experimentally as well as numerically.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.28 no.10
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• pp.1264-1270
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• 2004

In this paper research on micro solid propellant thruster is reported. Micro solid propellant thruster has four basic components; micro combustion chamber, micro nozzle, solid propellant and micro igniter. In this research igniter, solid propellant and combustion chamber are focused. Micro igniter was fabricated through typical micromachining and the effect of geometry was evaluated. The characteristic of solid propellant was investigated to observe burning characteristic and to obtain burning velocity. Change of thrust force and the amount of energy loss following scale down at micro combustion chamber were estimated by numerical simulation based on empirical data and through the calculation normalized specific impulses were compared to figure out the efficiency of combustion chamber.

• Proceedings of the KSME Conference
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• 2004.04a
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• pp.1280-1285
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• 2004

Microsystem technology has been applied to space technology and became one of the enabling technology by which low cost and high efficiency are achievable. Micro propulsion system is a key technology in the miniature satellite because micro satellite requires very small and precise thrust force for maneuvering and attitude control. In this paper research on micro solid propellant thruster is reported. Micro solid propellant thruster has four basic components; micro combustion chamber, micro nozzle, solid propellant and micro igniter. In this research igniter, solid propellant and combustion chamber are focused. Micro igniter was fabricated through typical micromachining and evaluated. The characteristic of solid propellant was investigated to observe burning characteristic and to obtain burning velocity. Change of thrust force and the amount of energy loss following scale down at micro combustion chamber were estimated by numerical simulation based on empirical data and through the calculation normalized specific impulses were compared to figure out the efficiency of combustion chamber.

• Journal of Aerospace System Engineering
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• v.18 no.3
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• pp.70-75
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• 2024

The insulation performance of a common bulkhead propellant tank for small launch vehicles with variations in insulation thickness was analyzed. The common bulkhead propellant tank composed of a single part allows for lightweight design, as it eliminates the need for tank connections. However, problems such as propellant loss and ignition delay due to heat transfer caused by temperature differences between oxidizer and fuel may arise. Therefore, it is essential to verify the insulation performance of the common bulkhead structure that separates the oxidizer tank and fuel tank. In this study, transient heat transfer analysis was conducted for propellant tanks with insulation thicknesses of (50, 55, 60, 65, and 70) mm to analyze the insulation performance using boil-off mass. Subsequently, the boil-off mass of the oxidizer generated during the first-stage flight time of the propellant tank was determined. The results confirmed that increasing the insulation thickness reduces the boil-off mass, thereby improving the insulation performance.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.968-971
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• 2017

In this study, manufacturing the Single Base propellant using a Nitroguanidine(NGD A%, NGD-B%, 2A=B). The Factors affecting the combustion rate are Moisture and Volatiles, Residual Solvents, Dimension. These Factors were analyzed and compared. Also, NGD-A% propellants and NGD-B% propellants were compared with firing test Test. As a result, NGD-A% was confirmed to have a higher speed at a similar pressure than NGD-B% propellants.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.2
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• pp.9-17
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• 2007

The performance test of recirculation line in propellant feeding system was carried out. Liquid oxygen was used as cryogenic propellant and helium was used as recirculation promotion gas. Tests were done in cases at atmospheric pressure and at pressure of 4 barg in the ullage space of propellant tank. Liquid oxygen recirculation flowrate with helium injection flowrate and temperature distribution along the line were measured. There was appropriate helium injection flowrate for gas-lift recirculation system. Test data were used to make calculation program by test data correlation method. In this paper the procedure of calculation was presented and the results were compared to test data.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.4
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• pp.73-78
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• 2020

The main reasons that ignite the propellant in the process of producing solid rocket motor are worker mistakes, wrong working process, mechanical defects, impact, friction, electrostatic and short circuits. In the past decades, many accidents have occurred in the process of producing solid rocket motor, accidents investigation have confirmed that the sensitivity of electrostatic is very high under specific condition. In this paper, we analyzed overseas accident cases and measured the sensitivity of electrostatic in the situation of confinement and pressure load by considering the manufacturing process. As a result of the test, the sensitivity of propellant was increased in the situation of confinement and pressure load and the propellant reacted more sensitively to electrostatic in the situation of confinement than pressure load.

• Aerospace Engineering and Technology
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• v.9 no.1
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• pp.133-142
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• 2010

Propulsion system of launch vehicle is composed with subsystems as propellant tank, pressurization system, propellant fill/drain system, valve operating system, purge system and so on. Among others, pressurization system is the most important subsystem, because of the real-time control part for pressure control of propellant tank. Therefore, it is the subsystem that must be primarily considered on conceptual design process. In this paper, the data of the previously developed pressurization systems were collected and the optimum configuration was selected by analysis of advantage and disadvantage of the systems.