• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.37-46
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• 2006

The outside environment is very severe while aircraft is cruising. Especially small particle of icing in cold air condition can have negative influence on aircraft performance. If ice particle is attached to leading edge of wing, it can change wing configuration and decrease flight quality. If icing particle is attached to inlet of engine, it can damage compressor blade and have negative influence to aircraft safety. We make icing simulation device with liquid air system for analyzing about variation of engine performance due to incoming of icing to engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1997.04a
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• pp.27-31
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• 1997

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.10
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• pp.56-60
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• 2006

It is essential to predict thermodynamic properties of combustion gas with respect to a propellant mixture ratio for the development of a gas generator for a liquid rocket engine. The present study shows the temperature measurement of exit combustion gas as a function of a mixture ratio through the series of combustion tests of a fuel-rich gas generator with liquid oxygen and Jet A-1. The measurements of dynamic and static pressures, and combustion gas temperatures allowed the estimation of thermodynamic properties like a specific heat ratio, a gas constant, and a constant pressure specific heat of the combustion gas. The comparison of the experimental results with predictions made by interpolation parameters obtained from the modification of the chemical equilibrium code indicates that the interpolation method calibrated using the temperature measurements can be utilized as an effective tool for the initial design of a fuel-rich gas generator.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.3
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• pp.54-65
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• 2022

We are developing a compact ignition device that can provide a multi-ignition capability for an upper stage methane engine of a two staged small satellite launch vehicle. Firstly, the multi-ignition device is designed and built as an integral part of an additively manufactured mixing head. Secondly, the ignition device requires no separate high-pressure vessels to store ignition propellants as they are branched out from the main feed lines for the mixing head. We performed experiments at various levels, including igniter autonomous tests, thrust chamber ignition and combustion tests on the new compact ignition device which is integrated in the thrust chamber of one-tonf class liquid oxygen/liquid methane engine, and confirmed stable ignition performance.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.34-37
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• 2009

The present study presents a numerical methodology for early conceptual trade-off study between propulsive performance, cooling efficiency, weight and size, in which combustion and cooling precesses in regeneratively cooled rocket thrust chamber are interactively simulated. To address the capabilities and reliability of the design tool, some application results are given involving contour design, performance analysis, and wall cooling prediction as well as a systematic design evaluation.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.5
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• pp.65-71
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• 2010

Performance test of the 75ton class turbopump turbine was performed. Using the measured turbine power characteristics in the wide-range operational conditions, variations of nozzle velocity ratio, total pressure loss, and relative flow angle to the pressure ratios and rotational speeds are quantified. Efficiency and nozzle exit pressure behavior was also investigated and compared with 30ton turbopump turbine data. A rotor blade was redesigned based on the test results and CFD analysis.

• Journal of the Korean Society of Propulsion Engineers
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• v.2 no.3
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• pp.90-98
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• 1998

The vaporization process of liquid oxygen(LOX) at high pressure environment is numerically investigated. The present vaporization model can account for the high-pressure effects such as ambient gas solubility, real gas behavior and variable properties. The predicted phase-equilibrium compositions for $N_2$/$H_2$ and $O_2$/He system are well agreed with experimental data. The LOX vaporization characteristics is parametrically studied for wide range of the operating conditions encountered in the high-pressure combustion process of liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.38-44
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• 2010

Performance test of the 75ton class turbopump turbine was performed. Through the turbine power measurement in the wide-range operational conditions, velocity ratio, total pressure loss, and relative flow angle characteristics was quantified. Efficiency and nozzle exit pressure behavior was also investigated and compared with 30ton turbopump turbine data. A rotor blade was redesigned based on the test results and CFD analysis.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.2
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• pp.1-7
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• 2001

To protect combustion chamber from high temperature combustion gas, regenerative cooling is used for most liquid rocket engine. Although regenerative cooling is the most effective way to protect the chamber from high heat flux, realization of this system requires detail analysis, manufacturing technique and high cost. To demonstrate the possibility of applying regenerative cooling to a real rocket engine, the hot fire test has been carried out for the sub-scale liquid rocket with the water cooling system. The main purpose of the test is to identify the problem area of design, safety and cost effective manufacturing technique. The coolant passage was 3 mm in width and wall thickness was 1 mm with stainless steel. Maximum combustion time and pressure were 60 seconds and 400 psi, respectively. The flow rate of coolant was reduced gradually from 2 kg/s to 0.12 kg/s throughout firing test, combustion chamber was visually examined and no dwfect was observed.

• Proceedings of the KSME Conference
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• 2007.05b
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• pp.3518-3523
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• 2007

An analysis program of specific impulse has been developed for a gas generator cycle rocket engine. The program has been verified by comparing the published performance data of the same cycle engine with RP-1 as fuel. A model for pressure drop of regenerative cooling and film cooling mass flow rate has been suggested to satisfy the necessary cooling condition with Jet-A1 as fuel. The engine mixture ratio is defined by the film cooling mass flow rate and the core mixture ratio. The optimal condition of the combustor pressure and engine mixture ratio has been found for maximum specific impulse.