• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.863-866
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• 2011

Recently environmental issue is more and more emphasized and 'Green Growth' became on of the key words of this Government. Based on this trend, the exhaust gases out of the gas generator cycle and the staged combustion cycle LRE whose propellants are kerosene and LOx were compared. For this purse, 8 tonf class of each cycle engine was designed and the amount and the components of the gases were investigated by using CEA. As expected, the staged combustion cycle engine generates less pollutants than the other cycle. In addition, the graphite that is generated by the gas generator can be reacted with the oxygen in the atmosphere creating additional pollutants.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.733-736
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• 2010

Korea Aerospace Research Institute (KARI) performed the preliminary design of liquid rocket engine high-altitude simulation firing test facility for the development and qualification of LRE for the 2nd stage of KSLV-II. The engine high-altitude simulation firing test facility, which are to be constructed at Goheung Space Center, will provide liquid oxygen and kerosene to enable the high-altitude simulation firing test of 2nd stage engine at ground test facility. The high-altitude environment is obtained using a supersonic diffuser operated by the self-ejecting jet from the liquid rocket engine.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.39 no.6
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• pp.485-496
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• 2015

The present study presents experimental results on the characteristics of emission spectra of kerosene/air swirl flames. The aviation fuel Jet A-1, which is used for the liquid rocket engines of the Korea Space Launch Vehicle, is used with three different swirlers to investigate the swirl strength effects. The emission spectra from the flames are measured with a spectrometer as the swirl strength and combustion air temperature are varied. Chemiluminescence intensities of $OH^*$, $CH^*$ and $C_2{^*}$ are identified from the spectra. The chemiluminescence intensities from the kerosene flames show sensitivity to the swirl strength and are affected by changes in the combustion air temperature. Among the three radicals of interest, $C_2{^*}$ show the most significant changes in chemiluminescence intensity with the swirl strength and equivalence ratio. The intensity ratios $I_{OH^*}/I_{CH^*}$ and $I_{C_2{^*}}/I_{CH^*}$ are adequate for indicating changes in the equivalence ratio with the air and fuel mass flow rates, respectively.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.6
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• pp.69-78
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• 2010

As part of preliminary study for development of 1,200 N-class bipropellant rocket engine with the concentrated hydrogen peroxide, bipropellant engine elements were designed and experimentally tested. The catalysts of $MnO_2$ and $MnO_2$ added Pb as an additive were compared to achieve high decomposition performance and the catalytic reactor with $MnO_2$ added Pb was designed and its decomposition efficiency of 97.2% was achieved. The autoignition tests of kerosene by decomposed hydrogen peroxide were carried out under various equivalence ratios to ignite without additional ignition sources. Autoignition were achieved in all experimental conditions and $C^*$ efficiencies at each condition were at or above 90%. From the measured thrust results, the highest value was 830 N which is in corresponds with 1,035 N at vacuum level assuming $C^*$ efficiency equals $I_{sp}$ efficiency.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.6
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• pp.56-62
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• 2011

An experimental study on design of a catalytic ignitor was performed to use an ignition source for a small bi-propellant liquid rocket engine which use hydrogen peroxide and kerosene as propellants. In the catalytic ignitor, hot gas of hydrogen peroxide which was decomposed by a catalyst induced autoignition of kerosene. Mass flow rate and O/F ratio for the ignitor were calculated by CEA code. A combustion chamber which had a quartz window and thermocouples was manufactured to determine whether the ignition is successful. Ignition performance was investigated according to exit area of fixed rings and mixture ratio. Results showed that reliable ignition performance was achieved at non-choking exit area of fixed ring and O/F ratio of 6~8.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.156-164
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• 2010

As part of preliminary study for development of 1,200 N-class bipropellant rocket engine with the concentrated hydrogen peroxide, bipropellant engine elements were designed and experimentally tested. The catalysts of $MnO_2$ and $MnO_2$ added Pb as an addictive were compared to achieve high decomposition performance and the catalytic reactor with $MnO_2$ added Pb was designed and its decomposition efficiency of 97.2% was achieved. The autoignition tests of kerosene by decomposed hydrogen peroxide were carried out under various equivalence ratios to ignite without additional ignition sources. Autoignition were achieved in all experimental conditions and $C^*$ efficiencies at each condition were at or above 90%. From the measured thrust results, the highest value was 830 N which is in corresponds with 1,035 N at vacuum level using 94.1% theoretical $I_{sp}$.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.10
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• pp.1038-1047
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• 2009

The development process of a gas generator for a 30-tonf pump-fed space liquid rocket engine is described. Starting from the development of an injector, followed by subscale and full-scale test specimens, the development of LOx/kerosene fuel-rich gas generator has been concluded successfully. Various analytical methods have been utilized in the course of design and the performance requirements have been verified experimentally through ignition tests, combustion performance and stability assessment tests and duration tests. The gas generator has proven its workability and stability within a defined operation window of varying chamber pressure and mixture ratio and demonstrated compliance to the performance and life time requirements.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.3
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• pp.62-80
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• 2021

The requirements for the next-generation propulsion system and for a good propellant have been summarized. The characteristics and effectiveness of kerosene, hydrogen, and methane, which are the fuels that are mainly attracting attention in Korea and abroad, were compared with each other. As a result of the comparison, methane was evaluated to be more advantageous than other fuels in reliability, cost, reusability, maintenance, eco-friendliness, safety, lifespan, technical difficulties, engine cycle selection, application of common bulkhead, and non-disassembly/reassembly delivery. And in terms of performance, the specific impulse of methane is higher than that of kerosene, so the efficiency of the launch vehicle can be increased. Methane's properties incluidng eco-friendliness, low-temperature combustion, long life, and maintenability make it beneficial for reuse and for the development of multi-purpose engines.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.196-200
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• 2006

This paper describes the current development status of the major subsystems, turbopump and gas generator, for a turbopump-fed liquid oxygen-kerosene rocket engine system. As a secondary stage of the liquid rocket engine development test, turbopump-gas generator powerpack tests are planned. The schematics of the test hardware and the test facility for the TP+GG coupled test are presented. The results of a preliminary analysis for operating regimes of the TP+GG coupled test are also presented.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.11
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• pp.1132-1138
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• 2008

A program that can simulate gas generator after burning liquid rocket engines was developed along with presenting the characteristics of the engines. The program was written in Matlab and used GUI interface so that many users can use it without any difficulties. The results of the program was compared with the real engine which was developed by the LRE advanced country. Most of the parameters concurred within 1% error expect for the pressure at the turbopump. The reasons of the large differences were supposed that pressure decreases at the schematics were smaller than that of the real engines.