• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.3
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• pp.41-48
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• 2008

One of main functions of the firing control system applied to a rocket propulsion test has been to provide electric current for ignition of a solid rocket motor. This paper describes the design and implementation of an enhanced firing control system for ground propulsion test that can also control and verify various types of squib events and flight sequences.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.130-133
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• 2009

The sequence of the propellant supply is very important for the reliable and safe operation of a LRE combustion test. So combustion performance tests were performed to find an optimum test sequence by changing supply time of propellants and purge gas in the moment of ignition and extinction. The liquid rocket engine consisted of a catalytic ignitor and six swirl-coaxial injectors which used hydrogen peroxide and kerosene. Conclusively, an optimum sequence was found for stable combustion in the moment of ignition and extinction.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.97-100
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• 2011

It is critical to set up the starting sequence of liquid rocket engines because not carefully arranged process can lead the engine damages. Thus, many efforts were made to prevent the hard start at the ignition. Hypergolic fuels are frequently used to ignite LRE and TEAL, one of the hypergolic fuel is also used for kerosene-LOx LRE ignition. However, since we are still lack of experiences igniting oxygen rich preburners of the staged combustion cycle engines, it would be helpful to estimate the TEAL ignition phenomena before the actual tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.750-755
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• 2011

The ignition characteristics of liquid rocket engine thrust chambers which have been developed by domestic technology were analyzed. Analysis results showed that low frequency fluctuation appeared in a partial ignition range according to different temperature profiles and vaporous state in the oxidizer manifold with startup sequences. This low frequency fluctuation wasn't developed as a malfunction factor, but this fluctuation is thought to be taken a continuous concern considering interfaces with engine system and launch vehicle.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.73-78
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• 2012

The ignition characteristics of liquid rocket engine thrust chambers which have been developed by domestic technology were analyzed through hot-firing tests. Thrust chambers used in hot-firing tests have different characteristics in terms of the injector for ignition, film cooling method and the position of the oxidizer inlet. Also, these thrust chambers used their respective startup sequences. Analysis results showed that according to temperature profiles of the oxidizer manifold, low frequency fluctuation was appeared in ignition area. This low frequency fluctuation didn't give rise to violent malfunction of the thrust chamber, but the continuous observation as a concern parameter in the side of interfaces with engine system and launch vehicle should be demanded.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.814-819
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• 2017

The objective of this paper is to observe effects of an ignition position and time on the ignition transition. A gaseous oxygen and liquid kerosene are used for propellants with the shear-coaxial injector. In order to study the ignition delay time and combustion instability intensity, the pressure transducer was used. The ignition position was changed with the injector spacer. Sequences except the igniter operation time were fixed to compare the ignition time only. Initial pressure peak and ignition delay time increased as the ignition time was delayed. Also, the unstable flame development zone was detected as the igniter was away from the injector.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1142-1143
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• 2017

In this paper, Underwater propulsion test of SWASH(Small Waterplane Area Single Hull) type underwater vehicle with hybrid rocket system is performed. Watertight structure is applied to prevent a combustion chamber from water, and the control logic is constructed by setting the watertight ignition sequence. As a results, It is confirmed that the ignition is stable in water, and the propulsion system works well along the configured control sequence.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.4
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• pp.85-90
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• 2018

The objective of this paper is to observe effects of ignition position and time on ignition transition. A gaseous oxygen and liquid kerosene mixture is used as propellant with a shear-coaxial injector. In order to study the ignition delay time and combustion instability intensity, the pressure transducer was used. Sequences, excepting igniter operation time, were fixed to compare the ignition time only. Initial pressure peak and ignition delay time increased as the ignition time was delayed. Additionally, an unstable flame development zone was detected when the igniter was away from the injector.

• Journal of the Korean Society of Combustion
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• v.5 no.1
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• pp.43-53
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• 2000

Experimental studies on determination of the supply leading time of propellants to combustion chamber have been made to stably and efficiently guarantee the ignition process with liquid rocket engine. The propellant used is a Jet A-1 as fuel and a liquid oxygen as oxidizer. Unlike impinging FOOF type of injectors are arranged radially and the designed O/F ratio is 2.34. The present experiment program also includes the stability on the quadlet type of ignitor using the triethylalumimum as an ignition source and injector life tests. Experimental results clarifies that the propellant supply through LOx leading to combustion chamber is proper for stable ignition and combustion processes based on the fuel and oxidizer manifold pressures, combustion chamber pressure, and the variation of flame length from the nozzle exit with lapse time, and shows that the leading supply time of propellants affects the engine performance little. The effect of positioning cooling holes is remarkable to protect the injector face.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.6
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• pp.11-18
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• 2014

The combustion flowfield at the near-injector region of a 50 N scale $N_2O/C_2H_5OH$ thruster was visualized using shadowgraph technique. The explosive ignition was occurred at the design spray condition, and the expanding combustion gas quenched the flame immediately. Approximately after 83 ms from the initial ignition, the propellant spray was re-ignited, and the flame was stabilized after 23 ms elapsed. In the increased oxidizer flow rate condition, the transient pressure at the moment of ignition was smoother than explosive ignition, and the blow down phenomenon was not appeared in the same operating sequence. In addition, the flame was stabilized within 17 ms, and it is caused by improved propellants mixing before ignition.