• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.10
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• pp.966-978
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• 2011

Transpiration cooling is the most effective cooling technique for the high-performance liquid rockets and air-breathing engines operating in aggressive environments with higher pressures and temperatures. When applying transpiration cooling, combustor liners and turbine blades/vanes are cooled by the coolant(air or fuel) passing through their porous walls and also the exit coolant acting as an insulating film. Practical implementation of the cooling technique has been hampered by the limitations of available porous materials. But advances in metal-joining techniques have led to the development of multi-laminate porous structures such as Lamilloy$^{(R)}$ fabricated from several diffusion-bonded, etched metal thin sheets. And also with the availability of lightweight, ceramic matrix composites(CMC), transpiration cooling now seems to be a promising technique for high-performance engine cooling. This paper reviews recent research activities of transpiration cooling and its applications to gas turbines, liquid rockets, and the engines for hypersonic vehicles.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.10
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• pp.66-72
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• 2003

The cooling mechanism for a liquid rocket engine of 10tf-thrust using kerosene as a fuel was studied from the viewpoint of both the regenerative and 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 en g i.ne could be improved.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.2
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• pp.29-37
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• 2015

This paper presents a numerical investigation on detonation of a kerosene-air mixture in the copper tube and the structural response associated with combustion instability in liquid rocket engine. A single step Arrehnius rate law and Johnson-Cook strength model are used to describe the chemical reaction of kerosene-air mixture detonation and the plastic deformation of the copper tube. The changes of flow field and tube stress which are induced by plastic deformation, are investigated on the different tube thicknesses and nozzle configurations.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.515-519
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• 2011

Structural tests for the mixing head of a 75tonf class thrust chamber were performed to verify structural stability. The mixing head of a thrust chamber is loaded by high pressure with regeneratively cooled fuel and cryogenic liquid oxygen(LOx) as well as it transfers thrust load generated by liquid rocket engine. Therefore structural stability of mixing head is a very important factor to work without any plastic deformation or structural failure. In this study, two mixing heads were manufactured using different welding methods, Tungsten Inert Gas(TIG) welding and Electron Beam Welding(EBW) and evaluated a structural stability. The results of structural tests showed that the mixing head assembled by EBW can withstand the applied design load without any structural failures and be structurally more stable than that of TIG welding.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.118-122
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• 2006

For the development of a gas generator of a liquid rocket engine, the prediction of thermodynamic properties of combustion gas with respect to a propellant mixture ratio becomes critical. The present study focuses on the temperature measurement of exit combustion gas as a function of a mixture ratio through combustion tests of a fuel-rich gas generator propelled by Lox/Jet A-1. The measurement of combustion dynamic and static pressures allowed indirect estimation of thermodynamic properties like specific heat ratio, gas constant, and constant pressure specific heat. Comparing the results with empirical prediction through an interpolation reveals that the interpolation method calibrated using temperature results can be utilized as an effective tool for the design of a fuel-rich gas generator.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.48 no.9
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• pp.717-724
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• 2020

We study the closing characteristics of a self-sustainable poppet valve which serves as a main oxidizer shut-off valve for liquid rocket engines. Numerical analysis for predicting closing transient responses are presented and the calculated results have been verified by a comparison with experimental data. The effective area of a pilot gas discharge system and the pressure distribution of passage flow around the valve moving part are shown to be main parameters in determining the closing characteristics for dry and cryogenic conditions, respectively. Moreover, it is presented that the passage flow pressure at the valve closing moment as well as the valve closing velocity can be effectively adjusted by the appropriate employment of the pilot gas.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.11
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• pp.1087-1093
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• 2008

Regenerative cooling passage to guarantee the thermal survivability in high performance rocket engine combustors could have complex configurations of the branching/merging of channels and flow turning, etc. By applying the classical hydraulic coefficients which can be found in the literature according to the flow conditions, hydraulic characteristics in regenerative cooling passages can be obtained effectively through dividing the pressure loss into friction loss and local resistance loss. Satisfactory agreement has been obtained by comparing the present results with experimental measurement of water flow test. In addition, the present results were in good agreement with CFD results when the actual coolant, kerosene was used. Therefore, the application of the present method is expected to be useful to design regeneratively cooled combustors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.594-596
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• 2010

Hot-firing tests were performed on a sub-scale gas generator for development of a 75 ton-class liquid rocket engine. This paper deals with the analysis results of low-frequency combustion instability that encountered during combustion tests of the gas generator.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.4
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• pp.84-90
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• 2004

A performance evaluation is made in terms of thrust, impulse bit. and specific impulses for a set of mono-propellant hydrazine thrusters producing 0.95 lbf of nominal thrust at an inlet pressure of 350 psia. With a brief description on the hot-firing test configuration and procedures. a typical data obtained from steady-state firing mode is given directly showing the variational behavior of propellant supply pressure, mass flow rate, vacuum condition, and thrust. The performance features are successfully compared to the reference criteria of 1-lbf standard mono-propellant rocket engine. Additionally. a statistical inter-thruster treatment is concisely depicted for the justification of selected thrusters as a grouped member of flight model for spacecraft propulsion system.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.3
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• pp.37-53
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• 2016

This paper deals with an optimal output control for Space Shuttle Main Engine (SSME), a liquid propellant rocket engine using a staged-combustion cycle. For this purpose, we modeled simplified mathematical model of SSME using each SSME component divided into 7 major categories and found trim points called Rated Propulsion Level (RPL). For design the closed-loop system of SSME, we designed optimal output feedback Linear Quadratic Regulation (LQR) control system using SSME linearized model under RPL 104% and demonstrated the performance of the controller through numerical simulation.