• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.6
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• pp.98-104
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• 2015

Excessive thrust acting on the rotor of pump can cause the damage of pump or the decrease of pump lifetime. Therefore, for ensuring the safety of LOX pump of a liquid rocket engine, the thrust of pump rotor is estimated by similarity tests. Axial thrust is indirectly measured by an axial thrust measurement unit positioned outside pump. Radial thrust is calculated based on pressure distribution of volute scroll. As a result, axial and radial thrust are increased when the flowrate of pump decreases. However, both thrusts do not affect the stability of pump rotor since their values are not large.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.2
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• pp.165-172
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• 2014

This study was focused on the investigation of the liquid propellants for the launch vehicle, past, present and future trends in worldwide. In general, storable and hypergolic propellants are used for the military purposes and comparing to this, the kerosine fuel and cryogenic propellants are used for the launching systems. Although liquid propulsion is seemed as a mature technology, the requirements of a renewed interest for space exploration has led to the development of a family of new engines, with more design margins, simpler to use and to produce associated with a wide variety of thrust and life requirements.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.5
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• pp.491-496
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• 2008

New concept ramjet propulsion system with liquid bipropellant rocket using "Green Propellant" hydrogen peroxide for launch stage is proposed. In this novel concept, hydrogen peroxide gas generator produces hot oxygen at launch stage and kerosene injects to this jet in combustor. For basic study of this new concept ramjet system, investigation of auto-ignition characteristics and combustion of decomposed hydrogen peroxide and kerosene was conducted. In various test cases, auto-ignition and stable combustion was verified. The combustion temperature of 400°C and Fuel/Oxidizer mixture ratio of 0.6 were the limit of auto ignition. Through the experiment results, the possibility of novel concept combined propulsion system using hydrogen peroxide gas generator is ascertained.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.9
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• pp.874-883
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• 2008

This study considers powered explicit guidance, one of the closed-loop guidance laws for satellite launch vehicles. The guidance algorithm employed here does not include the iterative procedure of the original algorithm. Also, the single-target algorithm to treat the general time-varying thrust profiles is described. The computer simulations for the 6-DOF launch vehicle model are performed to investigate the orbit injection accuracy of the guidance algorithm in the nominal/off-nominal flight conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.41-44
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• 2006

The propulsion system of VIA-50A airship consists of engine, generator, inverter, motor and propeller. The motor and propeller was designed that can be tilted to $120^{\circ}$ for thrust vector control. When the flight test was performed, various condition data of the airship were obtained by wireless telecommunication and analyzed in real-time. In this paper, we presented flight test results of propulsion system. Considering the designed requirement and normal range, we verified that all constituent part was operated in normal condition during the high altitude flight test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.66-68
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• 2011

Chollian is a geostationary satellite, and its bipropellant propulsion system is mainly composed of one main engine for orbit transfer and fourteen thrusters for on-station operations. The Chollian was launched successfully at Kourou Space Center in French Guiana. After it separated from the launcher, the propulsion system was initialised automatically. Then three times of main engine firing were successfully performed, and the target obit insertion was accomplished. This paper details the major CPS events during LEOP phase for the Chollian satellite.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.225-228
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• 2009

Hot-firing tests were performed on the gas generator which is a technology development/demonstration model for a 75 ton-class liquid rocket engine. A heat-sink type combustion chamber was used for initial performance examination of the injector and mixing head. This paper explains not only preparation works for hot-firing tests but also the acquired results such as pressure, temperature distribution, and pressure fluctuation.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.2
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• pp.56-61
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• 2011

Performance test of a full-scale turbopump assembly for a 75 ton class liquid rocket engine was carried out at full speed. Model fluid was used as a working medium: liquid nitrogen for the oxidizer pump, water for the fuel pump, and hot air for the turbine. The turbopump was operated stably, satisfying the performance requirements. Head coefficient and flow coefficient of the pumps remained constant at the speed-increasing period. In terms of performance characteristics of pumps and turbine, the results from the turbopump assembly test showed a good agreement with those from the turbopump component tests.

• Aerospace Engineering and Technology
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• v.11 no.1
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• pp.185-191
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• 2012

This paper developed Helicopter Chassis Dynamometer System(HCDS) to perform the bench test of the scaled rotor blade and to design a scaled model helicopter flight test bed and accomplished the scaled helicopter ground test. The scaled helicopter should be checked the relation of thrust and power input to maintain regular RPM by collective pitch angle versus throttle input. It showed hovering performance results of IGE with OGE, the max. F.M. was 0.76 without ground effect. The results of the chassis dynamometer test of scaled helicopter will usefully apply to design the scaled helicopter and evaluate the rotor blade performance.

• Aerospace Engineering and Technology
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• v.3 no.1
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• pp.197-204
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• 2004

A design of regenerative cooling system of 30 ton level thrust combustion chamber for ground test has been performed. The 1-D design code has been validated by comparing with the heat flux of the NAL calorimeter for high chamber pressure and water-cooling performance of the ECC engine of MOBIS. The present design code has been confirmed to predict accurately the heat flux and water-cooling performance for high chamber pressure condition. The maximum hot-gas-side wall temperature is predicted to be about 720 K without thermal barrier coating and the coolant-side wall temperature is less than the coking temperature of RP-1. The coolant temperature rises nearly 100 K with thermal barrier coating when Jet-A1 is used as coolant.