• 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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• 2005.11a
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• pp.293-298
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• 2005

Air-Launching is an effective method that can launch the 'Nanosat' with low launching cost. In this study, system and subsystem design of the air launching rocket for nanosats which perform a simple mission, have been performed. Foe this purpose, the WBS of the Air-launching Rocket System, and the subsystem schematics have been defined first. Based on these results, detailed configuration and DMU have been developed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.299-302
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• 2011

In the supersonic jet of a solid rocket motor, various noise is investigated. The purpose of this study is to attain and evaluating a design and manufacturing technique of the SRM noise reduction. In this study, the water is injected into the supersonic jet of the SRM to reduce the noise. Furthermore, the diffuser and stack are installed to suppress the SRM noise. Through the SRM ground tests, the noise is reduced approximately 20dBA with application of the diffuser/stack with water injection.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.125-129
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• 2009

The basic design of liquid rocket engine combustion chamber for a large space launch vehicle was described. It has vacuum thrust of 74.8 ton, vacuum specific impulse of 306.9 sec, chamber pressure of 60 bar, mass flow rate of 243.6 kg/s and combustion characteristic velocity of 1730 m/sec. The details of combustion performance and geometrical parameter were also given. The 75 ton combustion chamber consists of the combustor head with injector and the chamber/nozzle with regenerative cooling channels.

• 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.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.28 no.1
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• pp.16-23
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• 2004

Acoustic characteristics of unbaffled and baffled combustion chamber are experimentally investigated under atmospheric condition to preliminarily determine baffle for mitigation of combustion instability. To investigate the effect of the baffle which has several configurations such as radial baffles and hub/blade baffle, resonant-frequency shift and damping factors of the chamber were analyzed and compared quantitatively with those of the unbaffled combustion chamber. From a view of acoustic characteristics, radial baffles with several configurations have not much difference in resonant-frequency shift and damping factor ratio with each other. On the other hand, hub and blade baffle is very effective to suppress the first tangential mode which was found to be the most harmful acoustic mode in KSR(Korean Sounding Rocket)-III engine. But more study on design parameters such as hub size and axial length should be done for complete optimization of hub and blade baffle. The present study based on linear acoustic analysis is expected to be a useful confirming tool to predict acoustic field and design a passive control devices such as baffle and acoustic cavity.

• 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.

• 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.

• Journal of the Korea Institute of Military Science and Technology
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• v.18 no.1
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• pp.93-99
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• 2015

The dual pulse rocket motor(DPRM) distributes the propellant energy effectively via pulse separation device(PSD) to improve the range and terminal velocity of the missile. There are two types of PSD such as bulk head type and thermal barrier type. A subscale thermal barrier type DPRM was designed, manufactured, and tested. The results showed good understanding of the characteristics of the PSD and will be applied to the design of the full scale DPRM.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.748-751
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• 2010

To launch rocket on launch pad, propellants and gases are charged into the rocket by remote control system. Using pneumatic pressure-reducing regulators, kinds of gases with various pressure levels are supplied into launch pad. As most of operations for launching the vehicle are remotely controled in the launch control room, pressure pulsations due to rapidly gas supply at the upstream of regulators can make the required operating pressure range missed and cause damage to the regulators. In this paper, the optimum design methods of pressure regulators of pressure-reducing system on launch pad using high-pressure gases were investigated to solve the aforementioned problems and for stable gas supply to launch pad.