• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.8
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• pp.69-76
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• 2003

For stable combustion and safety of a structure of the propulsion system, a cooling system to the liquid rocket engine should be incorporated. In this study, Eulerian-Lagrangian scheme for two phase combustion, nongray radiation and soot formation effect, and film-wall interaction have been introduced to study the effect of film cooling. After briefly introducing the governing equation, combustion characteristics with change of wall temperature has been investigated by varying such parameters as fuel mass fraction for film cooling, diameter of the fuel droplet, overall mixture fraction of oxygen to fuel. Also, radiative heat flux is compared with the conductive one at the combustor wall.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.135-140
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• 2007

An analysis has been conducted of the optimal condition to maximize the specific impulse for a liquid rocket engine with film cooling. The present engine performance has been compared with the published conceptual design to be verified satisfactorily accurate. The optimal combination of film coolant flow rate and the regenerative cooling capacity has been found for maximum specific impulse. The optimal fuel pump pressure increases and the optimal film coolant flow decreases for a larger thrust engine. Higher turbine inlet temperature increases both the fuel pump pressure and the film coolant flow rate as the optimal condition. The coking temperature has the same qualitative effect as the turbine inlet temperature.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.16-22
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• 2012

Performance of a liquid rocket thrust chamber with regenerative cooling scheme has been numerically analyzed using in-house CFD code which can predict combustion/cooling performance and provide nozzle design parameters. This paper investigates trade-offs between combustion and cooling performance with varying amount of fuel directly injected into the chamber wall to form cooling films and mixture ratios for the peripheral injectors. Further efforts to verify/improve the simulation methodology including comparison with the firing test results are planned to make it a reliable tool to optimize the film cooling and other major design parameters.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.35-41
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• 2012

Performance of a liquid rocket thrust chamber with regenerative cooling scheme has been numerically analyzed using in-house CFD code which can predict combustion/cooling performance and provide nozzle design parameters. This paper investigates trade-offs between combustion and cooling performance with varying amount of fuel directly injected into the chamber wall to form cooling films. Also is analyzed the effect of varying mixture ratios for the peripheral injectors on combustion performance enhancement. Further efforts to verify/improve the simulation methodology including comparison with the firing test results are planned to make it a reliable tool to optimize the film cooling and other major design parameters.

• Aerospace Engineering and Technology
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• v.6 no.2
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• pp.141-149
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• 2007

It was carried out hot-firing test with a regenerative-cooled sub-scale combustor which was applied regenerative-cooling, film cooling and thermal barrier coating. Test results showed that cooling methods used in the combustor play an full role in the operation of the combustor under the design condition but it is occurred high frequency combustion instability due to unsteady flow of fuel by structural support ring inserted in fuel manifold. The flow pattern of fuel was improved by excluding the ring and it will be carried out additional hot-firing test to verify the combustion stability of modified combustor.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.83-88
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• 2003

An analytical study was carried out to evaluate the regenerative cooling characteristics in the liquid rocket engine of a 10tf-thrust level using kerosene as a fuel. As a supplementary cooling method, a radiative cooling was applied to the nozzle extension. It was found out from this work that the cooling system with the regenerative and radiation cooling only is not adaptable as a cooling method for the liquid rocket engine of a 10tf-thrust level using kerosene as a fuel for the 2nd stage of the space launch vehicle. So, additional cooling method, curtain cooling was introduced and analyzed. Curtain cooling was very effective to reduce the thermal and thermo-structural instability.

• Aerospace Engineering and Technology
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• v.8 no.2
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• pp.133-139
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• 2009

An analysis has been performed on the specific impulse for a liquid rocket engine of gas generator cycle. The present analysis method has been validated through the comparison of the optimal specific impulse for the 300t thrust conceptual engine against the published data. The engine specific impulse can be increased by applying film coolant decreasing the fuel pump head for regenerative cooling despite the decrease of specific impulse of the combustion chamber when the film coolant participates combustion more than the critical amount. The improved condition shows that higher combustion chamber pressure is achieved with less fuel pump head rise by additional film cooling.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.5
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• pp.78-84
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• 2007

In the regeneratively cooled combustion chambers of liquid rocket engine using hydrocarbon fuels, coking occurs as the wall temperature increases which results in compounds deposition on the wall of cooling channels. This phenomenon reduces cooling capability of the coolant which finally causes damage to the combustor by overheating of the chamber wall. In this paper, experiment results using an electrical heating equipment to simulate the regeneratively cooled channel are introduced and based on the results the compatibility of copper alloy with hydrocarbon fuel Jet A-1 is investigated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.78-82
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• 2003

The cooling mechanism for a regenerative cooling liquid rocket engine of 10tf-thrust using kerosene as a fuel was studied from the viewpoint of curtain cooling. Based on the concept of a highly-stratified gas flow in the combustion chamber, the cross section of the combustion chamber was spilt into 2 independent parts, core and exterior part. Additional fuel is injected into the exterior section and gas temperature can be reduced in the exterior section. Consequently, the heat flux into the coolant and wall temperature are reduced and the thermal stability of a liquid rocket engine could be improved.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.1
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• pp.20-28
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• 2010

The mixing characteristics of pylon injection in a Scramjet combustor and effects of film cooling to protect pylon from air-heating were investigated. Three-dimensional Navier-Stokes equations with $k-{\omega}$ SST turbulence model were used. Fuel hydrogen and air were considered as coolants. There were remarkable improvements of penetration and mixing rate with the pylon injection. There was also over-heating on the front surface of the pylon without film cooling. The coolant injected parallel to the front surface of the pylon protects the pylon from over-heating.