• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.2
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• pp.68-77
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• 2019

C/SiC composites were developed by a liquid silicon infiltration(LSI) method for use as heat-resistant parts of solid and liquid rocket propulsion engines. The heat resistance characteristics according to the composition ratio (carbon / silicon / silicon carbide) were evaluated by specimen test through arc plasma, supersonic torch test. An ablation equation for oxidation reactions was presented. Through the combustion test it was verified that various parts such as nozzle insert, exit cone and combustion chamber heat resistant parts for rocket propulsion can be manufactured and proved high ablation performance and thermal structure performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.2
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• pp.32-38
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• 2004

A comprehensive understanding is given for the hot-firing test results, which were obtained throughout the verification program of mono-propellant hydrazine rocket engines (thrusters) producing 0.95 lbf (4.2 N) of nominal steady-state thrust at an inlet pressure of 350 psia (2.41 Mpa). A scrutiny for the engine performance is made in terms of thrust and temperature behavior of steady state firing mode at the given propellant injection pressures: Pinj = 400, 250, 100, and 50 psi. The thrust and specific impulse are compared with a reference performance of 1-lbf standard rocket engines and their normalization procedure is introduced. A practical engineering approach to the data measurement and reduction is addressed, too.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.196-200
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• 2005

A theoretical model for the calculation of chemical equilibrium composition of propellant combustion product is briefly presented for the performance analysis of monopropellant hydrazine rocket engine. Analysis result is compared to that of test and evaluation of 1-lbf class thruster and is scrutinized primarily from the view point of ammonia dissociation fraction. Chemical equilibrium composition and average molecular weight is additionally depicted according to the variation of propellant inlet pressures and the varying nozzle area ratio. The theoretical analysis is tried as a way of derivation of design parameters for mid- and large-thrust class of monopropellant rocket engines.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.191-195
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• 2006

It is definitely required to control dispersion of the rocket engine performance in order to accomplish the mission of a launch vehicle successfully. A performance dispersion analysis was conducted for a gas generator cycle liquid rocket engine and the required pressure drops were estimated for engine tunning. As a result, the vacuum thrust dispersion of the engine was from +9.1% to -8.7% and the mixture ratio deviated from +9.7% to -9.6% from the nominal value due to the errors of components and the engine inlet condition of propellants. The required pressure drop in the LOx line to the combustor is higher than in the fuel line for same mixture ratio change.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.414-416
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• 2009

In order to build a transient simulation program for a high thrust liquid rocket engine(LRE), a static performance simulation program for components were made. The components were the thrust chamber (combustion chamber and supersonic nozzle), centrifugal pump (impeller and volute casing), impulse turbine, and flow control devices (control valve and orifice). Simplified mathematical models based on classical thermodynamic and inviscid theories were used to remove complexity and enhance the utility of the program. We examined the results of each program qualitatively for validate each component modeling.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.6
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• pp.91-97
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• 2011

A test range for a 75tf class gas generator cycle liquid propellant rocket engine is defined. The engine system test range is defined by the performance variation during flight, the dispersion after engine calibration, and additional margin. The component development test range includes the operation range corresponding to the engine system test range and the component performance margin.

• 한국전산유체공학회:학술대회논문집
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• 2008.03b
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• pp.316-319
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• 2008

A computational analysis has been conducted on the compressible flow in the turbine exhaust nozzle of the gas generator cycle liquid rocket engine. The commercial CFD code Fluent has been used. Four nozzle designs have been compared to select the turbine exhaust nozzle concept. Three candidates with single nozzle have comparable performance. The model with bifurcated nozzles shows significant performance loss. However it will be better in the view of balanced thrust distribution because of its symmetric geometry.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.122-129
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• 2008

Rocket Based Combined-Cycle(RBCC) engines are currently being explored as advanced propulsion for space transportation. JAXA has been conducting RBCC engine research by using various experimental facilities. In order to clarify the experimental results and contribute to the improvement of designing, the analysis of the RBCC engine in an ejector-jet mode was carried out using the CFD code developed in-house for unstructured grids. CFD replicated the characteristic flow structures. The numerical simulation of the pumping performance of the ejector driven by different rocket gases(He, $N_2,\;A_r$) and physical conditions were performed, and their effects on the performance were studied.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.51-56
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• 2011

A test range for a 75tf class gas generator cycle liquid propellant rocket engine is defined. The engine system test range is defined by the performance variation during flight, the dispersion after engine calibration, and additional margin. The component development test range includes the operation range corresponding to the engine system test range and the component performance margin.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.6
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• pp.10-18
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• 2015

Performance requirement analysis and weight estimation of a reusable launch vehicle with a rocket-based air-breathing engine(RBCC : Rocket Based Combined Cycle) were performed. Performance model for an RBCC engine was developed and integrated with flight trajectory model. The integrated engine-trajectory model was validated by comparing the results with those from previous research reference. Based on the new engine-trajectory model and previous research results, engine performance requirements were derived for an reusable launching vehicle with gross take-off weight of 15 tones. Dependence of the propellant amount requirement on the mode transition Mach number of the engine was also analyzed.