• Tribology and Lubricants
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• v.22 no.2
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• pp.105-111
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• 2006

The floating ring seal which is used in the high pressure turbo pump is frequently used in the oxidizer pump and the fuel pump of the turbo pump of the liquid propulsion rocket, because it is able to minimize clearance to decrease the leakage flow rate. Compared with contact seal, the floating ring seal has advantage of minimizing clearance without rubbing phenomenon. But, the floating ring seal has a tendency to increase instability in operating condition in the high speed region. In this research, we devised floating ring seal which is inserted bump in the outer surface in order to improve the stability in the high speed region. Through this work, we expect to improve stability of floating ring seal with increasing the direct damping coefficient of seal and decreasing the eccentricity ratio.

• Journal of Mechanical Science and Technology
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• v.20 no.6
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• pp.874-881
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• 2006

Combustion stability characteristics in actual full-scale combustion chamber of a rocket engine are investigated by experimental tests with the model (sub-scale) chamber. The present hot-fire tests adopt the combustion chamber with three configurations of triplet impinging-jet injectors such as F-O-O-F, F-O-F, and O-F-O configurations. Combustion stability bound-aries are obtained and presented by the parameters of combustion-chamber pressure and mixture (oxidizer/fuel) ratio. From the experimental tests, two instability regions are observed and the pressure oscillations have the similar patterns irrespective of injector configuration. But, the O-F-O injector configuration shows broader upper-instability region than the other configurations. To verify the instability mechanism for the lower and upper instability regions, air-purge acoustic test is conducted and the photograph or the flames is taken. As a result, it is found that the pressure oscillations in the two regions can be characterized by the first impinging point of hydraulic jets and pre-blowout combustion, respectively.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.5
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• pp.53-59
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• 2019

Solid fuel reacts with an oxidizer during combustion of a propellant to increase performance. Representative solid fuels are aluminum, cyclotrimethylenetrinitramine (RDX) and octahydro-1, 3,5,7-tetra nitro-1,3,5,7-tetrazocin (HMX). During combustion, these compounds generate white smoke by reacting with moisture and produce materials that are harmful to the environment, such as carbon monoxide, carbon dioxide, and methane gas. This study prepared a high-nitrogen-containing energetic material, hydrazinium 5-aminotetrazolate (HAT), which could be applied as a solid fuel. The compound was characterized by nuclear magnetic resonance (NMR) spectroscopy, and a thermal analysis was measured by differential scanning calorimetry (DSC). Also, the specific impulses and volumes of detonation gases were calculated using the EXPLO5 program.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.4
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• pp.98-106
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• 2000

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 ($\Delta$V) 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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• 2008.03a
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• pp.804-809
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• 2008

In this study, the cooling performance of a C/C composite material structure with metallic cooling tubes fixed by elastic force without chemical bonding was evaluated experimentally using combustion gas in a rocket combustor. The C/C composite chamber was covered by a stainless steel outer shell to maintain its airtightness. Gaseous hydrogen as a fuel and gaseous oxygen as an oxidizer were used for the heating test. The surface of these C/C composites was maintained below 1500 K when the combustion gas temperature was about 2900 K and heat flux to the combustion chamber wall was about 9 $MW/m^2$. No thermal damage was observed on the stainless steel tubes which were in contact with the C/C composite materials. Results of the heating test showed that such a metallic-tube-cooled C/C composite structure is able to control the surface temperature as a cooling structure(also as a heat exchanger), as well as indicating the possibility of reducing the amount of the coolant even if the thermal load to the engine is high. Thus, application of the metallic-tube-cooled C/C composite structure to reusable engines such as a rocket-ramjet combined cycle engine is expected.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.407-412
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• 2008

"KAGUYA"(SELENE) is a Japanese Lunar Explorer launched by H-IIA rocket from Tanegashima Space Center on 14 September 2007. The dual-mode bipropellant propulsion subsystem of KAGUYA includes two fuel tanks, an oxidizer tank, propellant and pressurant control components, twelve monopropellant 20N thrusters, eight monopropellant 1N thrusters, and a bipropellant 500N Orbit Maneuver Engine(OME). Once the KAGUYA separated from the rocket, it circled the Earth twice and traveled to the Moon, where it entered lunar orbit. All maneuvers were performed through multiple 500N OME/20N thruster firings. This paper describes the in-space performance of KAGUYA Lunar Explorer bipropellant propulsion subsystem.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.366-370
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• 2008

Mode analysis is very important for the development of liquid rocket engine in various applications. We developed a mode analysis program for the gas-generator cycle liquid rocket engine by proposing 13 independent equations with 13 independent variables which can be solved by Newton method. As an example we calculated the change of engine operating mode according to the control valve's loss coefficient change located in the gas-generator oxidizer supply line. And we concluded that this program can give basic idea for the mode analysis of gas-generator cycle liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.105-108
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• 2008

Nowadays, as there is so much interest in environment, hydrogen peroxide attracts attention as an eco-propellant. Hydrogen peroxide is widely used for mono-propellant of thruster, and oxidizer of bi-propellant rocket. Especially, it is used as mono-propellant of the thruster for attitude control of satellite and military weapons. So, the need of long time storage of hydrogen peroxide appears and storage test is required. In this paper, necessity of storage test of hydrogen peroxide and some conditions and methods are introduced. In addition, the results of storage tests under some condition are compared and analyzed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.251-254
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• 2008

Ground firing test facility and test field for firing test of hybrid rocket engine were constructed. Ground firing test facility were composed of hybrid rocket engine, thrust stand, oxidizer storage/supply system, control system and data acquisition system. Firing tests of thrust 50 kgf class were conducted. Stable performance data was obtained and operational reliability of ground firing test facility were found.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.11-15
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• 2012

A rotordynamic analysis is performed for a 7 ton class turbopump applied to the third stage LRE(Liquid Rocket Engine) of the KSLV(Korea Space Launch Vehicle). Based on the heritage of the developed experimental 30 ton class turbopump and developing 75 ton class turbopump for the KSLV first and second stage LRE, the 7 ton class turbopump is designed as an one-axis rotor turbopump. Two rotor systems comprised of one oxidizer pump assembly and the other fuel pump-turbine assembly are connected each other using a spline shaft and operating at a design speed. Through the rotordynamic analysis, it is investigated that the turbopump acquires sufficient separate margin of critical speed as a sub-critical rotor.