• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.11a
• /
• pp.863-866
• /
• 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
• /
• 2006.11a
• /
• pp.135-138
• /
• 2006

We developed an engine system design program by balancing the pressure-mass-power relation which can be acquired from each component's specification. In gas generator type open-cycle rocket engine system it is possible to distinguish the variables into two categories, which are input variables and requirement variables. We define 11 design variables corresponding to the 11 balance equations as functions of pressure, mass and power of target engine system. We solved these equations by Newton method. As an example we designed gas generator cycle engine system and finally we could conclude that this developed program is well suited to the engine system design.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.49-53
• /
• 2009

The paper presents the results of the experimental study about flow discharge characteristics of double swirl coaxial injectors for a liquid rocket engine. Flow discharge characteristics of injectors become one of critical design issues for LRE combustion devices. Tap water and liquid oxygen/kerosene were used for ambient and hot firing tests, respectively. A combustion discharge coefficient varies depending on a mixture ratio and a recess ratio, and magnitudes of the variations are different with respect to injector shapes and operating conditions. The variation of a combustion discharge coefficient with a LOx injector is considered to result from flame structure changes due to physical property changes.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.246-249
• /
• 2009

Test and launch facilities which use oxygen as the propellant of a launch vehicle have hazards of fire and explosion by the leakage of oxygen. Also, the personnel operating the facilities, which use the high-pressured gases like nitrogen and helium in the closed room, is exposed to the hazard of death from suffocation. Consequently, we should keep out of the hazards and the accidents by monitoring the contents of oxygen in the air. The method and the system construction for monitoring oxygen contents in the air and the results from its application to Naro space center are described on this paper.

• 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.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.17 no.5
• /
• pp.132-139
• /
• 2013

Liquid-propellant rocket engines are widely used all over the world, thanks to their high performances thrust, in particular high thrust-to-weight ratio. The sucess rate of the launching of the liquid propulsion is similar to the solid one even though it has more complex mechanical system. In general, liquid propulsion is seemed as a mature technology, the requirements of a renewed interest for space exploration has led to the development of a family of new engines, with more design margins, simpler to use and to produce associated with a wide variety of thrust and life requirements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.05a
• /
• pp.177-180
• /
• 2010

It has been studied the effect of mixture ratio and chamber pressure on variations of discharge coefficients. Combustion experiments of bi-liquid swirl coaxial injectors were conducted at fuel-rich conditions with liquid oxygen and kerosene. Using two types of injectors for the experiments, characteristics of the discharge coefficient have been identified from variations in a diameter of the fuel nozzle and a momentum ratio along with the change of a LOx spray angle. It is concluded that discharge coefficients do not vary because of no change of flame structures from the fact that the fuel swirl chamber is completely filled up with fuel flow.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.1016-1020
• /
• 2017

This study was conducted to develop a test facility that simulates the combustion instability that occurs in a real-scale liquid rocket combustor. A separate engine head with 3 injectors arranged in a row was designed/manufactured and verified through preliminary tests. The flow rate and spray pattern of the head were confirmed by the cold flow test. Next, propellant spray test and combustion test were carried out. A preliminary combustion test was carried out at 10 bar and the combustion chamber pressure was measured to be significantly lower than the target pressure. This is because it was a low pressure test, and it is expected to be resolved in the high pressure test in the future.

• 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.