• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.1
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• pp.64-70
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• 2007

A test facility was developed where an inducer for a liquid rocket engine turbopump can be tested using liquid nitrogen as a working fluid. At the facility, a hydrodynamic performance test and a cavitation performance test for an oxidizer turbopump were carried out. Head-flow relation at liquid nitrogen test was similar to the case at water test. However, cavitation performance at the liquid nitrogen was superior to the case at water test, which results from the thermodynamic effect of cavitation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.93-99
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• 2006

A test facility was developed where an inducer for a liquid rocket engine turbopump can be tested using liquid nitrogen as a working fluid. At the facility, a hydrodynamic performance test and a cavitation performance test for an oxidizer turbopump were carried out. Head-flow relation at liquid nitrogen test was similar to the case at water test. However, cavitation performance at the liquid nitrogen was superior to the case at water test, which results from the thermodynamic effect of cavitation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.155-158
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• 2007

As a secondary stage of the liquid rocket engine development test, turbopump-gas generator powerpack tests are being performed. The schematics of the test hardware and the test facility for the TP+GG coupled test are presented. The results of a preliminary test results for the verification of the propellant supply system of the test facility are also presented. Based on the preliminary tests results, the verification of the propellants supply systems of the facility system was performed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.288-293
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• 2006

The combustion performances and characteristics of the subscale liquid rocket combustion chamber are discussed in this paper. Subscale combustion chamber is composed of mixing head, ablative cooling cylinder, and water cooling nozzle. The mixing head has eighteen coaxial swirl injectors and one center coaxial swirl injector for ignition. The mixing heads employing the injectors of low different recess length are considered in this paper. The results of the firing test, comparison of performance, and characteristics of static and dynamic pressures of the four different mixing heads are described. The characteristics of combustion at design and of design points are also discussed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.156-160
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• 2006

This study investigated the characteristics of spray generated by a liquid-liquid co-axial swirl injector used in a combustor of the liquid rocket engine. The linear stability analysis[1] was introduced In liquid sheet breakup and Post[2]'s collision model which considers shattering was adopted on the collision model after breakup. Every model was implemented to KIVA[3], which was adopted as solve. To validate the implemented models the cases of high and low injection velocity were calculated respectively and each result agreed well with test results.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.177-180
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• 2009

A Possibility of combustion instability on longitudinal mode has a high level at large scale of L/D. Solid propellant has a metal particle and a grain of control to pressure oscillation. Solid rocket motor in slotted-tube grain controls pressure oscillation of longitudinal mode. If slot length is shot, pressure oscillation of longitudinal mode is amplified by cylinder part after middle phase of total burn time. A study has analyzed pressure oscillation of longitudinal mode at spectrum and acoustic modal analysis at pressure of result on static firing test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.78-81
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• 2009

A series of hydraulic tests of a fuel pump are performed using water at a room temperature. The pump is under development for 75-ton class liquid rocket engines of the open-loop gas generator type. According to the test results, the fuel pump satisfies its design requirement and its head and efficiency at the design flowrate are higher than the expected value by the computational analysis. Also, it is found that the pressure at the rear bearing outlet is higher than expected because the inlet of bypass pipe line is narrow. Furthermore, the flowrate of the secondary flow is estimated using the pressure difference of the elbow of the bypass pipe line.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.517-521
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• 2009

A unsteady internal ballistic analysis model was proposed to take account for the variance of local variance of pressure and velocity along the grain surface of a hybrid rocket combustor. The model of concern in the study was fairly comparable with the test result. The local variation of the oxidizer mass flow rate along grain surface results in chamber pressure, regression rate, and gas velocity change along its flow direction.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.1
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• pp.45-54
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• 2011

A program aiming at predicting dynamic characteristics of a Liquid Rocket Engine(LRE) was developed and examined to trace entire LRE operation. In the startup period, transient characteristics of the propellant flows were predicted and validated with hydraulic tests data. An arrangement of each component for the pipelines was based on an operating circuit of open cycle LRE. The flow rate ratio for the gas generator and the main chamber was determined to mimic that of real open cycle LRE. Individual component modeling at its transient was completed and was integrated into the system prediction program. Essential parameters of the component dynamic characteristics were examined in an integrated fashion.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.370-373
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• 2007

In this study, a hybrid rocket motor with separable and detachable oxidizer tank from combustion chamber is developed. Initially, the measured thrust of the motor showed about 30% of the design thrust since the oxidizer supply was not enough. In order to solve this problem, application is made to expand the orifice diameter of oxidizer injector empirically, so that the mass flow rate of oxidizer was improved. The improved performance was about 60% of design thrust, 18kgf, and thrust-to-weight ratio was reasonable, compared with other sounding rockets.