• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.05a
• /
• pp.187-192
• /
• 2006

The combustion performance tests of a 30 tonf-class full-scale combustion chamber performed with water as a coolant were described. The combustion chamber has chamber pressure of 53bara and propellant flow mass rate of 90kg/s. Since it was first firing test for 30tonf-class combustion chamber using channel cooling, water coolant mass flow .ate of 35kg/s and 18kg/s were performed which correspond to 110% of kerosene design volume flow rate and equivalent cooling performance of kerosene. The test results are described and the results showed that the water cooling performance of this combustion chamber is sufficient and the firing test is feasible using the kerosene as a coolant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.163-168
• /
• 2006

The combustion performance tests of a 30 tonf-class full-scale combustion chamber performed with kerosene as a coolant were described. The combustion chamber has chamber pressure of 53bara and propellant flow mass rate of 90kg/s. Since it was first firing test for 30tonf-class combustion chamber using kerosene cooling, kerosene coolant mass flow rate of 32kg/s which correspond to 120% of design mass flow rate were performed. Then, the firing test with kerosene mass flow rate of 25kg/s were successfully performed. The test results are described and the results showed that the kerosene cooling performance of this combustion chamber is sufficient and the firing test with regenerative cooling is feasible.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.10 no.4
• /
• pp.40-46
• /
• 2006

A 30-tonf-class fullscale thrust chamber for the application to a Low-Earth-Orbit Space Launch Vehicle has been combustion tested over the wide ranges of a mixture ratio and a chamber pressure. The thrust chamber designed for a pump-fed open cycle engine was tested with an ablative chamber instead of a regenerative one for the initial evaluation of its performance and function. The test results revealed stable combustion characteristics. The hardware survived the harsh environment and showed very sound functional characteristics. The measured combustion efficiency turned out to be 95% and a specific impulse at sea level was estimated as 254sec, which are comparable to or above the predetermined design values.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2007.04a
• /
• pp.309-312
• /
• 2007

The cryogenic fluid is the propellant for the liquid rocket engine. The design of space launcher vehicle is guided by minimum size and weight criteria, so the turbo pump solicits high impeller speed. Such high speed results in a zone of pressure drop below vapor pressure causing caivtation around inducer blades. The cryogenic fluid has different characters from isothermal fluid like water. The cryogenic fluid has very sensible thermodynamic properties and the phase change undergoes evaporative cooling. So, the developed code has to be modified cavitation modeling and it is added the energy equation for temperature sensitivity.

• 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
• /
• 2008.11a
• /
• pp.133-136
• /
• 2008

The present study presents experimental results of combustion tests of a fuel-rich gas generator associated with a turbopump. Five combustion tests had been successfully executed. Static pressures of the gas generator promptly reacted to propellant supply variations from the turbopump. A closed-loop test for driving the turbopump revealed no flaw. Exit gas temperature results are very similar to previous ones. An orifice was effective on the suppression of pressure fluctuations although tests conducted below 45 bar showed the same dynamic behaviour as that of component-only tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2012.05a
• /
• pp.489-492
• /
• 2012

Basic design result for propulsion system test complex(PSTC) of KSLV-II is briefly described. KSLV-II is a three stage launch vehicle will be used liquid rocket engine for each. PSTC is will be used for development and performances qualification tests of $1^{st}/2^{nd}/3^{rd}$ propulsion systems for KSLV-II. In the future, this result will be applied to critical design and manufacturing of PSTC after deciding lay-out and operating program.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.10a
• /
• pp.267-273
• /
• 2004

This paper targets to define controlling factors of pressure-coupled combustion response and estimate their effects on droplet evaporation process. Dynamic characteristics of hydrocarbon propellant vaporization perturbed by acoustic pressure are numerically simulated and analyzed. 1-D droplet model including phase equilibrium between two phases is applied and acoustic wave is expressed by harmonic function. Effects of various design factors and acoustic pressure on combustion response are investigated with parametric studies. Results show that driving frequency of acoustic perturbation and ambient pressure have important roles in determining magnitude and phase of combustion response. On the other hand, other parameters such as gas temperature, initial droplet size and temperature, and amplitude of acoustic wave cause only minor changes to magnitude of combustion response. Resultant changes in phase of heat of vaporization and thermal wave in droplet highly influence magnitude and phase of combustion response.

• Journal of the Korea Academia-Industrial cooperation Society
• /
• v.20 no.10
• /
• pp.100-106
• /
• 2019

Ground and high altitude simulated combustion experiments were conducted using a liquid rocket engine with hydrogen peroxide and kerosene as the propellant. A ground and high altitude simulated combustion test facility was constructed by installing a high altitude model diffuser and TMS (Thrust Measuring System) on a vertical combustion test bench. The thrust characteristics according to altitude were investigated using the combustion test equipment. The diffuser was designed on a 1:4.8 scale to verify the characteristics of the high diffusing diffuser and starting pressure. The cold flow tests were conducted using nitrogen gas, and the performance characteristics and starting characteristics of the scale down diffuser were verified. A diffuser and TMS were installed on the vertical combustion test bench, and the thrust correction equations for the system resistance were derived. The thrust correction equations were derived from the step test and vacuum step test before the actual hot firing test. Nozzles with an operating altitude of 10km were designed. Hot firing tests were conducted to analyze the thrust characteristics according to the operating altitude changes. The actual thrust was calculated using each correction equation with the thrust value measured by the TMS.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.24 no.1
• /
• pp.64-77
• /
• 2020

An electric-pump cycle, which is a propellant supply system for driving pumps of a liquid rocket engine using an electric motor, has the advantages of simple system configuration and easy control of supply flow rate and pressure. This paper investigates and analyzes the overseas research trends of the electric-pump cycle. In addition, the research and development country, performing organization, application, engine thrust, pump pressure increase, motor power, and rotation speed are summarized. Among them, the design variables of the overseas research that applied the upper-stage propulsion system with the thrust range of 0.445~2.2 kN could be used in the study of a similar electric-pump cycle in Korea.