• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.5
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• pp.107-114
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• 2018

Korea Space Launch Vehicle (KSLV)-I flight test data and the modified 1-dimensional steady state modeling data from the critical design results of the KSLV-II liquid oxygen filling system operation are compared to validate the reliability of critical design modeling. A comparison of major flow rates and pressure values between test data and calculation results are conducted. The relative errors relative to maximum total flow rate for each cooling, filling, and replenishment mode are determined within 6.7%. Calculated pressure values at the outlet of the pump and the inlet of flow control valves are within 5.1%. The pressure at the inlet of the launch vehicle for each operation mode are within the measured pressure range.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.3-8
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• 2009

Wall flush mounted injector with various orifice shape and injection conditions, were examined to enhance jet penetration and mixing in supersonic cross flow, in view of application to air-breathing accelerator vehicle. Orifice shapes with high aspect ratio were found to preferable for better penetration in the cold flow, and in the reacting flow for scramjet-mode combustion conditions. However, the effectiveness of the high aspect ratio was diminished in the dual-mode combustion conditions. Supersonic injection was applied to the high aspect ratio orifice, and further increase in penetration was observed in the cold and reactive flow for scramjet-mode combustion conditions, however, mixing enhancement due to mixing layer / pseudo-shock wave system interaction was dominant in the dual-mode combustion conditions. Difficulty in attaining ignition in the case with the high aspect ratio orifice was encountered during the combustion tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.93-97
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• 2010

The energy balance program which can balance the relations among energy, mass flow, pressure in the staged-combustion cycle of the liquid rocket engine has been developed. The modular approach has been chosen for the analysis; the engine cycle consists of the elements from the predefined component analysis program. The engine with the staged-combustion cycle has been decomposed into several principal component modules, such as a thruster chamber, turbopumps, turbines, supply system components and a pre-burner. The program has been verified with comparison of the results to the selected data of the space shuttle main engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.789-794
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• 2004

The pressure sensitive paint (PSP) technique has been well established in external flow field. However, there are still unresolved issues in internal flow field. This work was focused on the application to unsteady pressure measurement of fan flow field. The PSP measurement system was established and the image processing software was developed. First, the performance of PSP was investigated at the static cell. Then the unsteady PSP measurement was carried out at fan test facility. PSP data images were acquired from the suction and pressure surface of stator vanes. Pressure distributions on the surface of the stator vane were detected non-intrusively. The issues of image acquisition and image processing were clarified through the practical PSP application to fan flow field.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.573-576
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• 2004

KARI is achieving the KSLV program according to National Space Technology Development Program. In this paper, the authors are intended to introduce the Integrated Power Plant (abb. IPP) test facility which will be constructed for the variety of tests on KSLV program. IPP test facility refers to comprehensive testing equipment for liquid rocket launch vehicle. Using this facility, KARl can verify the adaptiveness of parts and subsystems for launch vehicle and finally can qualify the system characteristics of launch vehicle doing kinds of test including hot firing test. Using this facility, KARI can simulate the vehicle launching circumstances and it make to predict the performance of launch vehicle when its flight test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.301-306
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• 2008

A pyrogen igniter was designed to satisfy the required condition of kick motor system for the space launch vehicle. We analyzed the ignition characteristics and performed the combustion tests to verify the internal ballistic performance. In the design process, the arc-image test was carried out to find the sufficient heat flux as varying the initial pressure from 10 to 700kPa. The analysis indicated that the initial pressure condition would delay ignition time within a range from 100 to 500ms. The combustion test with an inert chamber was also performed to understand the ignition characteristics with the variation of the initial pressure of free chamber volume. Finally, we confirmed that the igniter could provide the acceptable energy to ignite the propellant of kick motor at the ground test. The result of the ground tests showed that the ignition delay time was within the design range at the atmospheric pressure condition.

• Progress in Superconductivity and Cryogenics
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• v.9 no.2
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• pp.31-34
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• 2007

Superconducting synchronous motors compared with conventional motors can reduce the motor size and enhance the motor efficiency for low-speed and high torque applications under the space constraints for propulsion system. Especially, homopolar superconducting synchronous motors (HSSMs) need less superconductor and lower magnetic flux density in superconductor field coil than air-cored superconducting synchronous motors (ASSMs). In addition, mechanical structure is more simplified and stability is increased because the superconductor field coil of HSSMs is not rotated in operation. In this paper, we present the outline of HSSMs including structure, characteristics and operational principles with the conceptual design of 5MW HSSM.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.654-657
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• 2011

For the successful development of korean exploraton program, KARI started development of ground test model for lunar lander from last year. In order to secure core technology for space propulsion system, Koreanization of propellant tank is proceeding and it will be used for final assembly and test for ground test model. In this paper, the development result of titanum tank shell and verification test result was presented.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.2
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• pp.60-71
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• 2017

A trajectory analysis program is developed using a 3DOF trajectory model for the performance analysis of geostationary launch vehicles by system options. Launch trajectory and the performance of injection at GTO was estimated using this program for several propellant options, engine types, number of engines and the location of launch site. Results of the analysis presents that the possibility of mission accomplishment by several design options using domestic launch sites and the development direction of GEO launch vehicles.

• Journal of Astronomy and Space Sciences
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• v.26 no.3
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• pp.403-416
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• 2009

COMS (Communication, Ocean and Meteorological Satellite) is a geostationary satellite and developed by KARl for communication, ocean and meteorological observations. It will be tested under vacuum and very low temperature conditions in order to verify thermal design of COMS. The test will be performed by using KARI large thermal vacuum chamber, which was developed by KARI, and the COMS will be the first flight satellite tested in this chamber. The purposes of thermal balance test are to correlate analytical model used for design evaluation and predicting temperatures, and to verify and adjust thermal control concept. KARI has plan to use heating plates to simulate space hot condition especially for radiator panels of satellite such as north and south panels. They will be controlled from 90 K to 273 K by circulating GN2 and LN2 alternatively according to the test phases, while the main shroud of the vacuum chamber will be under constant temperature, 90 K, during all thermal balance test. This paper presents thermal modelling including test chamber, heating plates and the satellite without solar array wing and Ka-band reflectors and discusses temperature prediction during thermal balance test.