• Journal of the Korean Society of Propulsion Engineers
• /
• v.20 no.4
• /
• pp.68-76
• /
• 2016

This research is carried out for the performance evaluation of the injector that is one of the critical components of bipropellant-rocket-engine. Spray characteristics are investigated in detail according to the recess length and injection pressure on the swirl-coaxial-injector using gaseous methane and liquid oxygen as propellants. A visualization is conducted by the Schlieren photography that is composed of a light source, concave mirrors, knife, and high-speed-camera. A hollow-cone-shape is identified in the liquid spray that is spread only by inner injector and the spray angle is decreased due to the diminution of swirl strength in accordance with the increase of the length of injector orifice. When the injector sprays the liquid through the inner injector with the aid of gas through the outer injector, the spray angle in external mixing region tends to increase with rise of the recess length, while in internal mixing region, it is decreased. It is also confirmed that the same tendency of the spray angle with recess length appears irrespective of the injection pressure of liquid spray.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.14 no.3
• /
• pp.39-51
• /
• 2010

Theoritical principles about the real gas equation of state are investigated and comparisons for the thermophysical properties of oxygen, hydrogen, and methane as the propellants of liquid rocket engine are carried out for the various conditions of pressure and temperature including supercritical conditions. The properties obtained using the real gas equation of state(Soave modification of Redlich-Kwong, Peng-Robinson equation of state, and extended corresponding states principle) have been compared with the results of applying the ideal gas equation of state. Differences of thermophysical properties among the models specifically at the liquid phase regime and their error ranges are addressed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.10a
• /
• pp.245-250
• /
• 2004

The objective of the present study is to conduct model combustion tests for various injectors to identify their combustion stability characteristics. Three different double swirl coaxial injectors with variation of a recess number have been tested for the comparative study of stability characteristic and flame structure. Gaseous oxygen and mixture of gaseous methane and propane have been employed for simulating actual propellants used for a fullscale thrust chamber. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.50 no.3
• /
• pp.165-172
• /
• 2022

Pintle injector is the most suitable injector for thrust control because it can control the area of propellant injection. Accordingly, the combustion test of multiple hole pintle injector and continuous type pintle injector was carried out in this paper using liquid oxygen and gas methane. The combustion performance of the two pintles was verified with the characteristic speed efficiency, and the experimental results were compared according to the O/F and combustion chamber pressure and under similar conditions. The efficiency of the multi hole pintle was generally somewhat higher than continuous pintle when pintle opening distance(the area of dispensing oxidizer) was in a 100% thrust condition.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.40 no.3
• /
• pp.222-229
• /
• 2012

The cooling of a injector effects on the vapor pressure of cryogenic oxidizer spray, and it decides the phase transition point at the ignition process, when the combustion chamber pressure increases drastically. The phase transition of oxidizer spray affects the ignition characteristics, and several ignition tests with the LOx/$GCH_4$ uni-element coaxial swirl injector was performed in the different initial temperatures of oxidizer injector, in order to investigate the effect of injector cooling on the ignition transient characteristics. At the transition point of oxidizer phase, where the combustion chamber pressure increased over the LOx vapor pressure, the temporary quenching phenomenon of the flame occurred. The lower temperature of chilled down injector and tubing tends to move up the phase transition earlier.

• Journal of Aerospace System Engineering
• /
• v.18 no.1
• /
• pp.53-63
• /
• 2024

Adopting an engine igniter with high efficiency and ignition performance is essential for reliable operation of liquid rocket engines. In this study, we developed a spark torch igniter for a 450 N-scale methane-oxygen liquid rocket engine by conducting numerical analyses, igniter manufacturing and validation. Specifically, we conducted a parametric study for maximizing the enthalpy at the igniter exit, specifically by adjusting the mass flow rate, nozzle area ratio, fuel-oxidizer mixture ratio, and the igniter length-to-diameter. The heat transferred via the igniter nozzle exit was computed using 3-dimensional numerical simulations. We also manufactured and tested the igniter based on a deduced design to confirm ignition performance of the designed spark torch igniter. The igniter developed through this study could contribute to the development of practical propulsion systems such as upper-stage engines of small launch vehicles.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.32 no.4
• /
• pp.61-66
• /
• 2004

The objective of the present study is to develop methodology for the assessment of combustion stability of liquid rocket injectors. To simulate actual combustion occurring inside of a thrust chamber, a fullscale injector has been employed in the study, which bums gaseous oxygen and mixture of methane and propane. The main idea of the experiment is that the mixing mechanism is considered as a dominant factor significantly affecting combustion instability in a fullscale thrust chamber. A single split triplet injector has been used with an open-end cylindrical combustion chamber. The characteristics revealed by excited dynamic pressures in gaseous combustion show degrees of relative acoustic damping depending on operating conditions. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.48 no.9
• /
• pp.703-709
• /
• 2020

Pintle injector is known to have been adopted as injector of Lunar Module Descent Engine (LMDE) and contributed to success of the Apollo program and recently used in merlin engine. In this study, 500N Lab-Scale pintle injector thruster was manufactured and the combustion experiment with LOx/GCH4 was conducted. However, the proto-type thruster was showed some problems, such as low combustion efficiency and melting of pintle tip. To solve these problems, the flow guide in pintle tip was suggested through the CFD simulation. After addition of flow guide module, the combustion efficiency increased and pintle tip did not melt until the end of combustion.