• Transactions of the Korean Society of Mechanical Engineers B
• /
• v.34 no.5
• /
• pp.563-568
• /
• 2010

Turbopump-gas generator open-loop coupled tests are performed during the development of a 30tonf-LOx/Kerosene rocket engine. In the turbopump-gas generator open-loop tests, the propellants to gas generator are supplied from the outlets of turbopump, while the gas exhausted from the gas generator is vented out to the atmosphere, instead of being used to turbine driving. This paper presents the objectives, procedure, and results of the open-loop coupled test, in addition to a schematic representation of the test apparatus and the operating conditions for the test facility system and control system. The results of turbopump-gas generator open-loop coupled test confirm chill-down procedure, startup characteristics, nominal operability and smooth shutdown of the open-loop coupled Test Plant in test conditions simulating engine system operation environment.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.05a
• /
• pp.159-162
• /
• 2008

In this study, various cycles of liquid rocket engines were surveyed and specifically gas generator after burning cycle was investigated for upper stage motors. The engines for the upper stage can be categorized into three group based on the cycles and propellants at the diagram. Kerosene engines which adapt the gas generator after burning cycle and are located in the region II, are characterized for high combustion pressure and complexity. This cycle usually needs more than two pumps to use the turbine power efficiently. The fuel line can be divided into the gas generator line and the combustor line, and only the gas generator line is need to be pressured more because the combustion pressure in the gas generator is much higher than that of the combustor. Basically, all the oxidizer goes into the gas generator and than to the combustor, thus the auxiliary LOx pump is not critically necessary. However, for the various reasons, the LOx line requires a booster pump. A gas generator after burning cycle engines produces relatively high specific impuls than that of the open cycle engines. Thus it is suitable for upper stages of launch vehicles.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.33 no.2
• /
• pp.92-97
• /
• 2005

A series of combustion test was done to verify the optimization result of a gas generator for a 10 ton thrust liquid rocket engine. An injector element is F-O-F impinging type injector and the test was conducted with kerosene/LOX propellants. Test results of combustion temperature and pressure show a very good agreement with optimal design result and verify that the design method was properly established. And turbulence ring revealed its effectiveness in enhancing combustion gas mixing and temperature difference in the radial direction showed only less than 15K. Also turbulence ring induced only 3.2% pressure loss in the combustion chamber, which is far less than conventional level observed in a gas turbine engine. Axial temperature distribution also shows that turbulence ring could effectively reduce about 10% or more in gas generator length if its location is properly selected.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.17 no.4
• /
• pp.81-88
• /
• 2013

It was designed and tested at the design point that an oxidizer rich preburner for a staged combustion liquid rocket engine propelled by kerosene and LOx. The oxidizer rich preburner was designed as some of LOx injected from the mixing head was burned with kerosene and the rest of LOx injected from injection holes in the regenerative cooling chamber was vaporized by combustion gas. The preburner is operated at OF ratio of 60 and combustion pressure of 20 MPa. The Preburner has a honey-comb type mixing head with simplex swirl injectors, a turbulence ring improving combustion stability and uniformity of product gas temperature distribution, and a nozzle simulating the duct. With the combustion test results at the design point, the oxidizer rich preburner showed high combustion stability and uniformity of product gas temperature distribution.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.26 no.2
• /
• pp.79-95
• /
• 2022

Since space exploration began in the 1950s, numerous upper stage engines have been developed and used based on various design concepts. In this paper, information of upper stage engines which developed or developing is analysed and their characteristics and performance are summarized. These days, there are many cases of commercial heavy launch vehicles applying upper stage engines using liquid hydrogen with expander cycle which launched recently. Engines operating by Kerosene seem to be close to its theoretical maximum performance based on past experiences. Meanwhile, engines using methane propellant, which has recently become an issue, are also undergoing many developments because of various advantages. Recently, private companies are actively participating in launch vehicle market, and there are many cases in which the government and companies jointly research of next-generation engine.

• 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.

• Journal of Aerospace System Engineering
• /
• v.11 no.2
• /
• pp.9-15
• /
• 2017

In order to verify the possiblity of improving the combustion performance of ethanol using zeolite catalyst and the characteristics of nitrogen oxides and carbon monoxide emission, micro gas turbine experiments were performed using catalytic reaction products, ethanol and kerosene as fuels and the results were compared. The thrust of the catalytic reaction product was lower than that of kerosene, but it was improved by 5% on average compared with the use of ethanol. Nitrogen oxides and carbon monoxide emissions of the catalytic reaction products were measured to be very low overall compared to kerosene. As a result, when the ethanol was reformed using the zeolite catalyst, the engine performance could be improved while maintaining the environment friendliness of the ethanol.

• Transactions of the Korean Society of Mechanical Engineers B
• /
• v.36 no.5
• /
• pp.563-569
• /
• 2012

An energy balance analysis is conducted for a 30 t thrust level liquid rocket engine. The relations between thrust and combustion pressure, between thrust and propellant flow rate, and between combustion pressure and fuel pump pressure rise are compared against those indicated by a published database of the existing rocket engines. A combustion pressure higher than the old design value is obtained, implying that the present design is high-performance oriented. The thrust to propellant flow rate ratio is the same as that of the existing engines, indicating that the specific impulse performance is at the usual level. The fuel pump pressure rise is found to be slightly high when the combustion pressure is considered, and it is attributed to the pressure budget of the present ground test engine not being optimized.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.8 no.4
• /
• pp.67-75
• /
• 2004

It is essential to develop a transient analysis model for the turbopump-fed type liquid rocket engine development, especially for deriving the number of test and its parameters. In this study we proposed a mathematical model of turbopump-fed type liquid rocket engine, and inspected transient mode changes of a rocket engine according to variations of thrust control valve opening ratio. To verify the results, we solved the same problem with AnaSyn software from Russia, and concluded that the results of transient code we developed deviated within 2% from AnaSyn results. Also, using the transient engine analysis code we showed the possibility to find out the system level design Parameters of the components. For example, we modeled a pressure stabilizer which is used to control the consistency of mixture ratio in the gas generator as forced damping system, and found the stability range of the natural frequency and the damping ratio with the transient engine system analysis code.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.31 no.4
• /
• pp.76-81
• /
• 2003

To investigate the convective heat transfer phenomena inside the Lox/Kerosene liquid rocket combustion chamber, hot fire tests were performed by using a water-cooled calorimetric chamber. The calorimetric chamber consists of one cylindrical section and nozzle section with independent cooling passage. To measure the heat flux, thermocouples were installed inlet and outlet of cooling passage of each section. The investigated range of combustion chamber pressure is from 100 psi to 300psi at fixed O/F ratio of 2.0 and radiation heat transfer from the hot gas to the surface is not considered. The measured heat flux was almost linearly depended on the chamber pressure.