• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2005.11a
• /
• pp.341-345
• /
• 2005

This paper is for the combustion characteristics of gas generator which drive 1.5MW-class turbo pump and runs in fuel-rich combustion regime with LOx/kerosene as propellant. The outline of development procedure of real scale high pressure gas generator is introduced and the relation between O/F ratio and outlet temperature and the molecular weight and specific heat ratio of combustion gas are described. The relation between O/F ratio and temperature is newly obtained at higher pressure and the molecular weight and specific heat ratio is modified and their validity is confirmed by the mass relation equation.

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

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.191-195
• /
• 2006

It is definitely required to control dispersion of the rocket engine performance in order to accomplish the mission of a launch vehicle successfully. A performance dispersion analysis was conducted for a gas generator cycle liquid rocket engine and the required pressure drops were estimated for engine tunning. As a result, the vacuum thrust dispersion of the engine was from +9.1% to -8.7% and the mixture ratio deviated from +9.7% to -9.6% from the nominal value due to the errors of components and the engine inlet condition of propellants. The required pressure drop in the LOx line to the combustor is higher than in the fuel line for same mixture ratio change.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2007.11a
• /
• pp.119-122
• /
• 2007

The present study employs experimental approach to identify combustion stability characteristics of fuel-rich gas generators. The gas generator of interest, fueled by LOx and Jet A-1, experienced combustion instability coupled to a longitudinal resonant mode of the combustion chamber at about 1200 Hz. The occurrence of instability is strongly associated with acoustic boundary condition at the exit .and axial location of maximum heat release. As a result, stretching heat release zone in the axial direction by increase of the fuel nozzle diameter has dramatically stabilized combustion.

• Journal of the Korean Society of Combustion
• /
• v.5 no.1
• /
• pp.43-53
• /
• 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 ILASS-Korea
• /
• v.18 no.1
• /
• pp.27-34
• /
• 2013

A preburner is one of the key components for a staged combustion cycle engine fueled by kerosene and Lox. Since it has oxygen rich combustion inside, temperature control is very crucial. The temperature of the exhaust gas should be low enough not to burn turbine blade and yet high to keep the efficiency high. In addition temporal and spatial deviations also managed strictly. Conventionally, the required average and maximum temperature are determined by engine system and the preburner should be developed to meet the criteria. Currently being developed preburner has 50K spatial temperature deviation requirement. It was estimated by numerical simulations and proven by tests. The numerical analysis were done with both supercritical condition and normal conditions. The tests results showed that the temperature deviations were less than expected, and the results from the test and simulations were well agreed when the supercritical conditions were considered. Above all, since the gas temperature created by the preburner is very stable with minimum deviation, the preburner developed can be used to drive a turbine and for gas-liquid combustion chambers.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.9-12
• /
• 2010

This paper discusses experimental data for the assessment of flow discharging characteristics of double swirl coaxial injectors operating at fuel-rich conditions. Combustion tests employing liquid oxygen and kerosene (Jet A-1) were conducted and a discharge coefficient was utilized for defining flow characteristics. A mass flow rate, a pressure, and a temperature were measured to estimate discharge coefficients. Fuel injectors revealed a fixed value of a discharge coefficient regardless of matched LOx injector design, chamber pressure, and mixture ratio. However, oxidizer injectors showed varying discharging coefficients depending on chamber pressure and mixture ratio. Flame structure variations seem to affect flow discharging characteristics of the oxidizer side.

• The Bulletin of The Korean Astronomical Society
• /
• v.37 no.2
• /
• pp.148.1-148.1
• /
• 2012

로켓 혹은 우주발사체의 주엔진에는 대부분 연료와 산화제를 연소시켜 나오는 에너지를 사용하는 화학로켓이 주종을 이루어 왔다. 이러한 로켓엔진에서 그동안 연료로는 수소계, 탄화수소계, 아민계 등 다양한 화학물질이 사용되어 왔으나, 산화제로는 강한 산화성을 나타내면서 밀도가 높은 몇몇 물질만이 제한적으로 사용되어져 왔으며, 최근에는 주로 액체산소(LOx)와 사산화질소(N2O4)가 사용되고 있다. 그러나 산화제 중 액체산소는 극저온이면서 상대적으로 밀도가 낮고, 사산화질소는 강한 독성을 지니고 있으며 액체로 존재하는 구간이 좁아 연구 목적의 소형발사체를 구현하는 것에는 많은 어려움이 있다. 이러한 이유로 최근 소형발사체 개발분야에서는 상온저장성이면서 친환경적인 과산화수소(H2O2)와 아산화질소(N2O)를 산화제로 활용하는 것에 대한 관심이 고조되고 있으나, 대형 추진기관을 개발하는 연구자들로부터는 액체산소를 사용할 때 보다 엔진 자체의 비추력이 상대적으로 낮다는 이유로 활용이 외면되어 온 것이 사실이다. 본 연구에서는 엔진 자체의 추진성능 보다는 사실상 발사체의 목적이라고 할 수 있는 추진단 속도증분을 성능의 지표로 삼아 평가하였으며, 결과를 통하여 과산화수소와 아산화질소의 높은 밀도가 엔진의 낮은 비추력을 충분히 보상할 수 있음을 보였다. 과산화수소와 아산화질소는 교육/연구용 소형발사체 구성에 충분히 활용가능한 산화제이며, 실제 발사에서 충분한 비행성능을 기대할 수 있는 물질로 평가할 수 있다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.1144-1147
• /
• 2017

It was reviewed about the pressure oscillation on the inlet of turbo-pump at the moment of engine startup and shutdown. Specially, This research was performed how much is the effect of PSD(Pogo Suppression Device) about the pressure oscillation on the inlet of turbo-pump at the moment of engine startup and shutdown. For analysis, propellant tank PSD and Engine are modelled with Flowmaster which is the commercial 1D program. As the analysis results, even though the PSD is installed in the pipeline, the pressure drop or rising at the moment of engine startup and shutdown is same compared to the case without PSD. However, it was confirmed that PSD reduces the pressure oscillation of the high frequency band as the original purpose of PSD.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.598-602
• /
• 2017

Korea Aerospace Research Institute has been developing 3-stage launcher KSLV-II, which can inject 1.5-ton satellite into sun synchronous orbit (SSO). For development process, Test Launch Vehicle(TLV) adopting the $2^{nd}$ and $3^{rd}$ stage of KSLV-II will be scheduled to launch in 2018. The propulsion system of TLV is composed of $2^{nd}$ stage engine system (ground type) and propellant supply system including LOX, Kerosene tanks. Until now, system integration of engineering model of TLV and delivery of qualification model have been done. In this paper, the product assurance activities for propulsion system KSLV-II will be illustrated.