• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.125-129
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• 2009

The basic design of liquid rocket engine combustion chamber for a large space launch vehicle was described. It has vacuum thrust of 74.8 ton, vacuum specific impulse of 306.9 sec, chamber pressure of 60 bar, mass flow rate of 243.6 kg/s and combustion characteristic velocity of 1730 m/sec. The details of combustion performance and geometrical parameter were also given. The 75 ton combustion chamber consists of the combustor head with injector and the chamber/nozzle with regenerative cooling channels.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.165-169
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• 2010

Liquid Rocket Engine is generally composed of extremely low and high temperature field. So that the component works properly including the electric component, the heat insulator should be applied appropriately. There are three steps. First, the heat source components should be defined and temperature field analyzed. Second, the heat transfer of pipes between the heat sources should be analyzed. Third, the components and pipes before and after applying the heat insulator should be analyzed. Finally, the optimized heat insulator depth can be calculated. In this paper, the procedure of this steps is established and investigated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.185-188
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• 2007

The main objective of this research effort was to develope the performance of C/SiC composites manufactured by LSI (Liquid Silicon Infiltration) method for solid and liquid rocket propulsion system and ensure the performance analysis technique. The various carbon preform were manufactured by filament winding, tape rolling, involute layup and stack molding process. For the best performance of thermal and mechanical properties, many process conditions were tested and selected by varying preform, the content of SiC, temperature, impregnation resin and chemical vapour reaction. In conclusion, the high performance and reliability of C/SiC composite were proved for solid and liquid rocket propulsion system. And the performance analysis technique related to mathematical ablation model was originated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.928-929
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• 2017

A 75 ton class liquid rocket engine has been developed for KSLV II. Ground firing tests of the rocket engine were conducted to characterize and quantify its structural vibration under combustion testing environments. In this study, evaluation of dynamic response characteristics of the engine system was presented to verify its structural integrity and structural design during ground firing tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.235-240
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• 2010

Transpiration cooling is the most effective cooling technique for liquid rocket and air-breathing engines operating in aggressive environments with higher pressures and temperatures. Combustor liners and turbine vanes are cooled by the coolant(air or fuel) passing through their porous walls and also the exit coolant acting as an insulating film. However, its practical implementation has been hampered by the limitations of available porous materials. The search for more practical methods of increasing the internal heat transfer within the walls has led to the development of multi-laminate porous structures, such as Lamilloy$^{(R)}$ and Transply$^{(R)}$. This paper reviews recent research activities of transpiration cooling for the propulsions of liquid rocket, gas turbine, and scramjet.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.15-18
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• 2009

Technology demonstration for the development of a 75-tonf liquid rocket engine(LRE) thrust chamber for a space launch vehicle has been started on the basis of the previously acquired 30-tonf LRE technologies. For this purpose, a technology demonstration plan was established upon considering the currently available firing test facility in Korea and performance evaluation firing tests were performed on technology demonstration model thrust chambers under a restricted test condition. This paper describes the plan and current status of technology demonstration for a 75-tonf LRE thrust chamber.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.85-92
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• 2006

A design program has been developed to predict film cooling performance of a liquid rocket engine. A thermal protecting effect of low mixture ratio gas layer has been analysed by CFD. A one-dimensional film cooling model based on the CFD results has been implemented to the previously developed design program of regenerative cooling. Satisfactory agreement has been achieved by comparing the predicted maximum heat flux at the throat of a subscale chamber and the average measured value, and the predicted nozzle average heat flux and the measured value for a full scale chamber with film cooling. It is ascertained that the film cooling is effective to reduce the throat heat flux in rocket engine chamber.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.1
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• pp.83-89
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• 2011

A design plan was proposed for determining combustor configuration of regenerative- cooled liquid rocket engine in the process of preliminary design. Rocket performance and regenerative cooling results were calculated using the properties of combustion gas estimated in CEA. For required thrust, chamber pressure, atmosphere pressure and propellant mixture ratio the mass flow rate of propellants and combustor performance were predicted by one-dimensional and experimental correlations. Finally, determinable plan for the contour of combustor were presented through Rao nozzle design method.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.9
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• pp.781-788
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• 2012

Flow in the oxidizer manifold of a liquid rocket combustor has been analysed using an open source CFD toolbox, OpenFOAM. The applicability of OpenFOAM to the problems with complex geometries involving porous media zones for simulating the pressure drop induced by the injectors has been evaluated by performing turbulent, incompressible steady-state flow analysis. The usefulness and applicable area of the OpenFOAM as a design evaluation and analysis tool will be confirmed and enlarged by further evaluation with various computational cases representing major physical phenomena in rocket combustion devices.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.37-42
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• 2010

A design plan was proposed for determining combustor configuration of regenerative- cooled liquid rocket engine in the process of preliminary design. Rocket performance and regenerative cooling results were calculated using the properties of combustion gas estimated in CEA. For required thrust, chamber pressure, atmosphere pressure and propellant mixture ratio the mass flow rate of propellants and combustor performance were predicted using one-dimensional and experimental equations. Finally, determinable plan for contour of combustor were presented through Rao nozzle design method.