• Journal of the Korean Society of Propulsion Engineers
• /
• v.20 no.3
• /
• pp.96-102
• /
• 2016

The 7tonf-class power pack test at turbopump test facility in Naro space center was performed for confirmations of starting/running/ending operation characteristics before 7tonf rocket engine hot-firing test. The dynamic pressure mounted on a combustion chamber of gas generator is measured under 0.2 bar which does not conditioned to the unstable combustion. The analysis results of RPM and acceleration sensors mounted on the turbopump, the power pack test was performed to the estimated RPM with the stable combustion.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.05a
• /
• pp.25-28
• /
• 2010

The Cold flow and ignition tests have been performed for a technology demonstration model of 75-tonf liquid rocket engine thrust chamber which was designed and manufactured on the basis of the previous development experience of a 30-tonf liquid rocket engine thrust chamber. The hydrodynamic characteristics of the facility supply pipelines and the filling time of the cooling kerosene were obtained through the cold flow tests. The ignition cyclogram was determinded using the results and the ignition test was successfully carried out. The acquired data and test technique of present ignition test will be used in hot firing tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.97-100
• /
• 2009

Cold flow and ignition tests were performed for a technology demonstration model of a 75-tonf thrust chamber which is a candidate liquid rocket engine for a next Korea Space Launch Vehicle. The test facility was modified to support the new concepts of the thrust chamber such as ignition system, film cooling and LOx leading supply. The hydrodynamic characteristics of the supply pipelines, thrust chamber and igniter as well as the filling time of the propellants were obtained through the cold flow tests on the LOx and kerosene and the ignition cyclogram was determined using the results. The ignition test was successfully accomplished according to the cyclogram and therefore, a basic information was obtained for further hot firing tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2007.11a
• /
• pp.170-173
• /
• 2007

Cooling characteristics of a liner were investigated by a film cooling method using a gas nitrogen in a rocket engine. High temperature gas of this test was made by mixing liquid nitrogen with combustion gas of a liquid rocket. A supply system of gas nitrogen was additionally constructed to the existing test facility of liquid rocket engine, and a new test section consisted of a liner and a gas injection ring was manufactured. A 10 second firing test for finding cooling characteristics of the liner was successfully conducted and liner surface temperatures and hot gas temperature was obtained.

• Journal of the Korea Academia-Industrial cooperation Society
• /
• v.20 no.10
• /
• pp.100-106
• /
• 2019

Ground and high altitude simulated combustion experiments were conducted using a liquid rocket engine with hydrogen peroxide and kerosene as the propellant. A ground and high altitude simulated combustion test facility was constructed by installing a high altitude model diffuser and TMS (Thrust Measuring System) on a vertical combustion test bench. The thrust characteristics according to altitude were investigated using the combustion test equipment. The diffuser was designed on a 1:4.8 scale to verify the characteristics of the high diffusing diffuser and starting pressure. The cold flow tests were conducted using nitrogen gas, and the performance characteristics and starting characteristics of the scale down diffuser were verified. A diffuser and TMS were installed on the vertical combustion test bench, and the thrust correction equations for the system resistance were derived. The thrust correction equations were derived from the step test and vacuum step test before the actual hot firing test. Nozzles with an operating altitude of 10km were designed. Hot firing tests were conducted to analyze the thrust characteristics according to the operating altitude changes. The actual thrust was calculated using each correction equation with the thrust value measured by the TMS.