• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.10
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• pp.102-111
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• 2004

The procedure of design and manufacture of sub-scale combustor using bipropellant swirl injector with external mixing for a liquid rocket engine are described. The results of cold flow test, ignition test and combustion test of the sub-scale combustor are also given in this paper. The sub-scale combustor uses liquid oxygen(LOx) and kerosene as propellants and has a injector head, an ablative material combustor wall and a water cooled nozzle. The injector head has LOx manifold, fuel manifold, fire face plate, one center swirl injector and 18 main swirl injectors. The cold flow, ignition and combustion tests were successfully performed without damage of combustor. Results of hot firing tests show that combustion efficiency meets the target of design and operations of start and stop cyclogram are stable and high frequency combustion instability does not occur.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.38 no.12
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• pp.1435-1440
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• 2014

The effects of the specimen size and temperature on the compressive strength of ATJ graphite were investigated. Compressive tests were conducted in accordance with ASTM C 965 at room temperature, $700^{\circ}C$ and $900^{\circ}C$. Three types of cylindrical specimen at room temperature were used in uniaxial tests, where the diameter-to - length ratios were one to two for the ASTM standard specimen, one to one for the Type I specimen, and one to 0.5 for the Type II specimen. Two kinds of cylindrical specimens, with and without antioxidant coating, were tested at elevated temperature. The Compressive strength of the expanded specimens(Type I, II) were slightly higher than that of standard specimen at room temperature. The compressive strength of a specimen with antioxidant coating increased as the temperature increased to $900^{\circ}C$. In contrast, that of the non-coated specimen decreases sharply due to the oxidation of the specimen.

• Journal of the Korean Society of Propulsion Engineers
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• v.2 no.3
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• pp.54-61
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• 1998

A numerical simulation is attempted for the regenerative cooling heat transfer processes of the liquid propellant rocket engine. The heat transfer from the combustion gases to the thrust chamber wall is called gas side heat transfer. This heat is conducted radially to the coolant through the carbon deposit and metallic wall of thrust chamber Finally, this heat is convected away by the coolant flowing along the passages in the thrust chamber. The equivalence of these three heat fluxes of the above processes is utilized to determine the coolant side wall temperature, gas side wall temperature and the heat flux. When the number and shape(width, height) of coolant passages, the shape(size) of thrust chamber, oxidant and fuel properties, coolant properties, oxidant/fuel mixture ratio, coolant inlet temperature, the thickness of carbon deposit formed along the thrust chamber wall during combustion are given, reasonable radial direction temperature distributions and heat fluxes along the thrust chamber axis are obtained.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.1-9
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• 2020

Using kerosene and liquid oxygen, 1.5-tonf class liquid-liquid pintle injector with rectangular two-row orifice was designed and manufactured. The combustion test of the pintle injector was carried out to verify the combustion performance and combustion stability under a supercritical condition which is the actual operation condition of the liquid rocket engine. The combustion test result showed that the pintle tip was damaged by the high temperature combustion gas in the high-mixed ratio recirculation zone of the combustion chamber. To solve this problem, the insert nozzle was installed in the pintle injector to increase cooling performance at the pintle tip. As a result of the hot firing test, installation of the insert nozzle, AR and BF had a great effect on pintle tip cooling performance.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.503-506
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• 2009

In this study, we presented the results of static firing tests on the PE/LN2O hybrid rocket motor, which has a thrust of 650 kgf level. Through the early tests, we found that the combustion chamber pressure and the thrust were lower than design values because an actual oxidizer flow rate was less than that expected. In order to complement this result, the methods of decrease of nozzle throat and the increase of oxidizer mass flow rate were conducted in the next experiment, and we studied the combustion phenomena with the experimental results. Also we compared and analyzed a difference of combustion characteristics on scale effect. It show that a sub-scale motor regression rate was a little less than that of a lab-scale motor with the same oxidizer mass flux. Results of this study might be used as a basic data for development of hybrid sounding rocket.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.8
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• pp.684-691
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• 2015

A sudden increase in combustion pressure is observed in the ducted rocket combustion test equipped with pipe shaped and converging nozzle exhaust tubes. This study aims to understand the physical mechanism of abrupt change in combustion pressure using thermal choking in the exhaust tube. Results confirmed that the thermal choking of the flow inside the exhaust tube was responsible for the sudden increase in combustion pressure. Also, high pressure exponent of solid propellants is critical sensitive to the occurrence of thermal choking exhaust pipe. Additionally, numerical simulation showed that the sudden increase in combustion pressure was less possible in diverging pipe because thermal choking is more reluctant to occur.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.5
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• pp.27-36
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• 2013

Linear stability analysis for combustion instability within a cylindrical port of solid rocket motor has been conducted. The analysis of acoustic energy has been performed by a commercial COMSOL code to obtain the mode function associated to each acoustic mode prior to the calculation of stability alpha. An instability diagnosis based on the linear stability analysis of Culick is performed where special interests have been focused on 5 stability factors(alpha) such as pressure coupling, nozzle damping, particle damping and additionally, flow turning effect and viscous damping to take into account the flow and viscosity effect near the fuel surface. The instability decay characteristics depending on the particle size is also analyzed.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.5
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• pp.19-26
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• 2019

Paraffin fuels usually have low combustion efficiency due to discharged unburnt droplets from the nozzle. Therefore, optimization of the post-chamber is becoming an important factor for performance. In this study, combustion experiments were conducted by changing either the length or diameter of the post-chamber to reveal the combustion behavior of paraffin fuel for hybrid rocket. As a result, the combustion efficiency improved due to the increase of the residence time as the post-chamber length increased. On the other hand, it is found that the influence of the diameter change was not significant compared with the case of variable post-chamber length.

• Journal of Aerospace System Engineering
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• v.14 no.5
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• pp.66-72
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• 2020

In this study, a method for increasing the initial water supply was employed to protect the water injection nozzle by the flame when supplying the water to the steam generator. During the initial steam generator test, the flow rate was controlled by using the only venturi, but cooling water was not supplied to the combustion chamber at the beginning of combustion, thereby resulting in damage to the water nozzle. To solve this problem, a venturi and an orifice were configured in parallel to increase the initial supply flow rate to form a differential pressure between the water manifold and the combustion chamber. Venturi and orifice supply sequences were established through the water flow tests, and combustion tests were conducted for final verification. Consequently, a continuous supply of the cooling water at the beginning of combustion was achieved, and the experiment was successfully performed without damaging the cooling water nozzle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.247-250
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• 2005

In the case of appling a unielement injector developed for a full scale liquid rocket combustor, a operating condition or configuration of the injector is changed by combustion pressure, arrangement and injector quantity of a full scale liquid rocket combustor. In order to verify application, swirl coaxial injectors propelled by jet-A1 and liquid oxygen are tested at different conditions of a combustion pressure, a flowrate and an injector length. As a test result, the application of the present swirl coaxial injectors is excellent because an efficiency of a characteristic velocity is increased at the each test condition beyond that variation of dynamic pressure intensity is small.