• Transactions of the Korean Society of Automotive Engineers
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• v.3 no.1
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• pp.45-53
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• 1995

스파크 점화기관에서 화염전파과정에 관한 연구를 수행하기 위하여는, 초기화염핵 구간에서의 화염의 형성과 발달의 거동을 정확히 파악하여야 한다. 그러므로 화염핵의 형성과 발달에 영향을 미치는 최소 화염핵 크기의 이론적인 계산을 수행하였다. 이론식을 정립하기 위하여 열점화 이론을 이용하였다. 최소 화염핵 크기를 계산하기 위해 열전도 계수, 화염온도, 층류연소속도, 기타 열역학적 상태량 등을 계산하였다. 계산에 의존한 화염핵 크기의 신뢰성을 확인하기 위하여, 점화에너지를 변화시킬 수 있는 점화장치를 사용하여 실기 운전을 통하여 희박연소 한계가 그 때의 화염핵이 성정할 수 있는 영역이라고 가정하여 그 정확도를 확인 하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.869-872
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• 2011

Ignition tests of an oxidizer rich preburner for a staged combustion cycle liquid rocket engine were performed to evaluate combustion performance. Design operation conditions of the tested oxidizer rich preburner are about 60 of OF ratio and 20 MPa of combustion pressure. The entire kerosene and some LOx injected into the mixing head is burned in combustion chamber and the remaining LOx injected through center holes of combustion chamber is vaporized. Full flow ignition method with hypergolic fuel was used. Each propellant was supplied in two stages for soft ignition. Test results, low frequency oscillation was occurred in low flow rate conditions under 45% of design flow rate. Stable ignition in the course of design combustion pressure was able to induce by minimization of low flow rate ignition region to escape low frequency oscillation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.705-706
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• 2010

The design analysis for the ignition transient combustion characteristics of a Kick Motor igniter indicated that the initial pressure condition would delay ignition time within a range from 100 to 500 ms. In the development tests, we confirmed that the igniter could provide the acceptable energy to ignite the main propellant at ignition transient.

• 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.

• 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.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1999.10a
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• pp.29-29
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• 1999

점화천이(Ignition transient)는 일반적으로 점화 신호를 주는 순간부터 로켓모타가 평형상태(equilibrium) 혹은 설계 작동조건에 도달했을 때까지의 시간 구간으로 정의되며, 이 시간 동안에 높은 연소실 압력 증가율에 의한 동연소효과(dynamic burning effect)와 연소가스의 높은 cross-flow 속도에 의한 침식연소효과(erosive burning effect)에 의해 추진제의 연소증가 현상을 일으킨다. 이런 두가지의 증가된 연소효과와 더불어 과대하게 설계된 점화기로부터 유입되는 질량유량에 의해 연소실을 채우는 시간(Chamber filling) 중에 압력 과잉(pressure overshoot)이 나타난다 그러나 동연소효과 및 침식연소효과의 경우 추진제의 종류와 로켓모타의 형상 등에 의해서 복합적으로 나타나는 현상이기 때문에 모든 로켓모타에 대해서 공통으로 적용할 수식이나 방정식이 존재하지 않는다.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.3
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• pp.77-81
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• 2012

In a cryogenic propellant tank the pressurant is contracted due to heat loss and the propellant itself evaporates. On a restartable propulsion system such phenomena are more intensive because the propellant contacts with the pressurant on the larger surface during the coast flight. Such heat and mass transfer phenomena should be considered for estimating the amount of pressurant. On the hypothesis that the heat and mass transfer quasi-equilibrium is achieved during the coast flight, the calculation process of the equilibrium pressure is presented. On the process the amount of loaded helium on the Falcon-1 second stage is calculated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.201-205
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• 2011

In a cryogenic propellant tank the pressurant is contracted due to heat loss and the propellant itself evaporates. On a restartable propulsion system such phenomena are more intensive because the propellant contacts with the pressurant on the larger surface during the coast flight. Such heat and mass transfer phenomena should be considered for estimating the amount of pressurant. On the hypothesis that the heat and mass transfer quasi-equilibrium is achieved during the coast flight, the calculation process of the equilibrium pressure is presented. On the process the amount of loaded helium on the Falcon-1 second stage is calculated.