• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.149-152
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• 2008

Effects of chamber configuration on combustion characteristic velocity of full-scale combustion chamber for 30-tonf-class liquid rocket engine were studied. The configurations of combustion chamber are ablative and channel cooling chamber (${\varepsilon}$=3.2) which have detachable mixing head, and single body regenerative cooling chamber which has nozzle expansion ratio of 3.5 and 12, respectively. The combustion chambers have chamber pressure of 53${\sim}$60 bar and propellant mass flow rate of 89 kg/s, and the injectors of all combustion chamber have recess number 1.0 and double-swirl characteristics. The hot firing test results at design point show that the combustion characteristic velocity of the regenerative cooling chamber which has nozzle expansion ratio of 12 is higher than that of other combustion chambers. The reasons for the above result are the increases of combustion pressure and enthalpy of kerosene which is heated due to cooling of the chamber wall before injection into the combustion field.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.134-138
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• 2005

The structural analysis and water pressure test of regenerative liquid rocket thrust chamber cooling channel specimens are performed at room temperature. material properties of copper alloy are obtained by uniaxial tension test at room temperature and used of elastic-plastic structural analysis. The plate type of cooling channel specimen are manufactured and performed water pressure test in order to confirm the analysis results. The differences between results of elastic-plastic analysis and that of water pressure test of cooling channel specimen are small and find that manufacturing process affect the structural stability of cooling channel very much because cooling channel thickness is small

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.194-197
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• 2007

The dome stretching and tension test have been made to obtain a forming limit curve(FLC) for the copper alloy which is used for manufacturing the regenerative cooling chamber. For experimental survey of the forming limit curve, we have used in-plane tension specimen to obtain tension-compression strain state and also out of plane specimen to obtain tension-tension strain state through dome stretching test. All specimens are divided into longitudinal and radial direction specimens by the manufacturing method. The test results shows that in tension-tension region, copper alloy possesses a maximum major strain of 62.3% and maximum minor strain of 58.6%. In the tension-compression region, maximum major strain is 60.5% and maximum minor strain is 25.8%.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.31-34
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• 2007

Performance tests of supersonic exhaust diffuser were conducted by using hot combustion gas for simulating high-altitude environment. The test rig consists of a combustion chamber, a vacuum chamber, water cooling ring and diffuser. Before combustion experiments, the preliminary leak tests were carried out on the liquid rocket engine and diffuser by using high pressure nitrogen(30barg) and a vacuum pump. The leak test results showed that there was no leaks at high pressure and vacuum pressure conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.2
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• pp.46-56
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• 2013

The test facility with hot-air supply system is required to develop transpiration cooling materials and experimentally evaluate its performance. In the study, the facility consists of an arc-plasma generator, plenum chamber suppling cold air, and test section was designed and an internal flow analysis was executed. From CFD results, it was confirmed that the designed plenum chamber thermally safeties and ideally mixes with plasma gas and cold air in the chamber. In addition, validity of design for supplying homogeneous flow to the test section was confirmed by this analysis.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.807-812
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• 2011

Fuel-side cold-flow tests were performed on the technology demonstration model of a 75 ton-class liquid rocket engine combustion chamber for the first stage of the Korea space launch vehicle II. Pressure drop in the cooling channels of the combustion chamber was measured by changing fuel mass flow rate through a pressure regulating system. Pressure drop in each segment of the chamber could be obtained and a lot of pressure drop was caused by high flow velocity in the nozzle throat segment. The accuracy of a hydraulic analysis method for calculating a pressure loss in cooling channels could be verified by applying it to the cold-flow test conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.56-61
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• 2012

Fuel-side cold-flow tests were performed on the technology demonstration model of a 75 ton-class liquid rocket engine combustion chamber for the first stage of the Korea space launch vehicle II. Pressure drop in the cooling channels of the combustion chamber was measured by changing fuel mass flow rate through a pressure regulating system. Pressure drop in each segment of the chamber could be obtained and a lot of pressure drop was caused by high flow velocity in the nozzle throat segment. The accuracy of a hydraulic analysis method for calculating a pressure loss in cooling channels could be verified by applying it to the cold-flow test conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.3
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• pp.84-91
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• 2014

The flow rate control of a cavitating venturi has been investigated with downstream pressure variation. A set of cavitating venturies for a liquid rocket engine thrust chamber firing test facility have been designed and manufactured. The flow characteristics of the cavitating venturies have been analyzed by experimental and computational methods. Results showed that constant mass flow rate condition was established by the cavitation inside the venturi. However, upstream pressure less than the actual design pressure of the cavitating venturi could not supply a constant flow rate.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.1
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• pp.81-84
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• 2004

A water-cooled calorimeter chamber with 8 cooling channels based on KSR-III sub-scale chamber has been designed and manufactured. One dimensional empirical correlation has been used at the design stage and full three-dimensional CFD analysis has been conducted to confirm the cooling condition for hot fire test is safe. Predicted heat flux is accurate around the nozzle throat when the thermal resistance of carbon deposit is considered. However relatively large difference is observed in chamber part.