• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.468-471
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• 2012

This paper describes a construction scheme of hot backup or triple modular redundancy control system in a ground hot-firing test facility to carry out performance assessment of propulsion system used in a space launch vehicle. It was possible for a hot backup redundancy control system with manual operated console to simulate TMR control system. A console layout of control system in control center to restrict imprudent works of operators was proposed.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.9-16
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• 2020

A performance evaluation of the 200 N-class GCH₄/LOx small rocket engine was performed through ground hot-firing test. As a result, the combustion pressure and thrust raised with the increase of the oxidizer supply pressure, and thus the specific impulse, characteristic velocity, and their efficiency increased. The characteristic velocity was measured at about 90% performance efficiency. The change of chamber aspect ratio did not affect the performance of the rocket engine in the test condition specified. In addition, uncertainty evaluation was conducted to ensure the reliability of the test results.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.4
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• pp.94-102
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• 2011

Sub-system requirements of a pressure-fed hot-firing test facility for performance assessment of a Liquid Rocket Engine(LRE) thrust chamber using Liquid oxygen and kerosene were described. These requirements were based on the experience of construction and operation of the ground hot-firing test facility which was used for the development of the KSR-III and a 30 tonf-class LRE thrust chamber. So it is expected that this paper is used as a basic material and an itemized previous review statement for the design and construction of a large hot-firing test facility.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.63-71
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• 2010

Sub-system requirements of a pressure-fed hot-firing test facility for performance assessment of a Liquid Rocket Engine(LRE) thrust chamber using Liquid oxygen and kerosene were described. These requirements were based on the experience of construction and operation of the ground hot-firing test facility which was used for the development of the KSR-III and a 30 $ton_f$-class LRE thrust chamber. So it is expected that this paper is used as a basic material and an itemized previous review statement for the design and construction of a large hot-firing test facility.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.29-33
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• 2009

On the basis of the development experience of a gas generator for the 30 ton-f thrust liquid rocket engine combustor a Subscale Ground Firing Test Facility was designed and fabricated for a gas generator for the 75 ton-f thrust liquid rocket engine combustor. The Subscale Ground Firing Test Facility developed is going to be used to develop 75 ton-f class gas generator. Acquired data and test technique from this facility will be used to develope the high performance liquid rocket engine combustor and the Ground Firing Test Facility. This report describes the improved Subscale Ground Firing Test Facility for 75 ton-f class gas generator and results of the cold flow test.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.109-117
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• 2006

Hot firing performance test of hydrazine decomposition catalyst used for monopropellant thruster of satellite and launch vehicle was performed on the ground. A test equipment for hot firing performance measurement of catalyst test was developed in collaboration with Hanwha Corp., and the catalyst firing performance were tested with the equipment. After a reaction delay time, a catalyst activity and a granule stability were measured for 2 times, satisfactory results were obtained such as 25msec, 2%, $704^{\circ}C$ for each test items on the average. In addition, the current development status of domestic prototype catalyst and its decomposition performance test results are presented.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.299-304
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• 2005

The design procedures of full-scale combustion chamber with chamber pressure of 53bara, mass flow rate of 90kg/s, combustion efficiency of $94\%$ and specific impulse at ground of 253sec were described. The details of combustion performance and geometrical parameters were also given. Full-scale combustion chamber consists of the combustor head with injector/baffle and the chamber/nozzle with regenerative cooling channels. The design results of combustion chamber with ablative materials, detachable injector head with SUS baffle or baffle injector and chamber body for ground hot firing tests were given in this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.85-92
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• 2020

Ground hot-firing tests were conducted to analyze the combustion performance according to the characteristic lengths 1.37 m, 1.71 m, and 2.06 m of the combustion chamber in 200 N-class GCH4-LOx small rocket engine. Thrust, specific impulse, and characteristic velocity at the steady-state could be obtained as the key performance parameters of the rocket engine. The performance characteristics acquired through the test were compared and analyzed with the theoretical performance calculated from CEA analysis. Observation of the influence of characteristic length on the combustion performance indicates that an optimal characteristic length shall remain between 1.71 m and 2.06 m.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.2
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• pp.144-149
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• 2014

A ground firing test for hot-fire performance evaluation according to the characteristic length($L^*$) variation of thrust chamber was carried out for the hydrazine thruster which may be employed in space launch vehicles. A scrutiny into the performance characteristics of each thruster is made in terms of thrust, specific impulse, response characteristics, and characteristic velocity at steady-state firing mode with propellant inlet pressure of 2.41 MPa (350 psia). Through the test results, it has been verified that performance of characteristic velocity and specific impulse degrades as the characteristic length deviates from that of the standard model. Thus, it is confirmed that the thrust chamber configuration of standard model was suitably designed for the requirement specified.

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