• Journal of the Korean Society of Propulsion Engineers
• /
• v.15 no.2
• /
• pp.68-73
• /
• 2011

For the launch preparation of KSLV-I, gaseous nitrogen with various level of pressure and cryogenic liquid nitrogen are required. Nitrogen Supply System on launch complex has been developed to perform the production of high pressure gaseous nitrogen, the production of gaseous nitrogen with temperature of 273 ${\pm}$ 2K for protection purge of launch vehicle after loading of propellant and the supply of cryogenic liquid nitrogen for cooling of fuel (kerosene) and oxidizer (liquid oxygen). The operational instability of vaporizer mainly caused by its heat transfer characteristics which sensitively depends on the atmospheric conditions was removed by introducing parallel installation of two vaporizer and their switching operation. The developed Nitrogen Supply System carried out its function successfully in preparation of KSLV-I flight tests.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.14 no.1
• /
• pp.56-64
• /
• 2010

A system design has been conducted of the liquid rocket engine for Korean launch vehicle (KSLV-II, Korea Space Launch Vehicle II). The present turbopump-fed liquid rocket engine of vacuum thrust 76 ton and vacuum specific impulse 297 sec adopts gas generator cycle. The combustion pressure of the regeneratively cooled combustor is 60 bar. The propellant is LOx/kerosene. The engine is started by pyrostarter and the combustor is ignited by TEA (TriEthylAluminium). The engine system performance and the subsystems performance requirements are given through energy balance analysis. The combustion pressure, specific impulse and the engine mass are analyzed to be reasonable comparing with the published data. The startup analysis method which will be used in the future has been validated against the turbopump-gas generator coupled test. The tuning method for performance variation of the engine which is not actively controled has been prepared by mode analysis and performance deviation analysis.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.800-803
• /
• 2010

Tetrahydrotricyclopentadiene(below THTCPD) isomer is a good candidate for the high performance liquid fuel component because of high density and heat of combustion value. But it has bad fluidity. The object of this study was to find out the optimal reaction condition to improve the fluidity of THTCPD because initial reactant is solid state, So we have carried out isomerization reactions using aluminum chloride catalyst under varying the reaction condition such as reaction temperature and reaction solvents. The results showed that when using aluminum chloride catalyst, isomerization reaction was more active in dichloromethane(methylene chloride: MC), 1,2-dichloroethane(ethylene chloride: EC) and chloroform than n-hexnae and toluene and was effected by reaction temperature.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.752-757
• /
• 2010

For the launch preparation of KSLV-I, gaseous nitrogen with various level of pressure and cryogenic liquid nitrogen are required. Nitrogen Supply System on launch complex has been developed to perform the production of high pressure gaseous nitrogen, the production of gaseous nitrogen with temperature of $273{\pm}2K$ for protection purge of launch vehicle after loading of propellant and the supply of cryogenic liquid nitrogen for cooling of fuel (kerosene) and oxidizer (liquid oxygen). The operational instability of vaporizer mainly caused by its heat transfer characteristics which sensitively depends on the atmospheric conditions was removed by introducing parallel installation of two vaporizer and their switching operation. The developed Nitrogen Supply System carried out its function successfully in preparation of KSLV-I flight tests.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.21 no.3
• /
• pp.61-67
• /
• 2017

In this study, Burning Area Analysis Program (BAAP) was developed by using VOF method to estimate the burning area of 3D shaped grain. The parametric study of mesh size, burning rate and time interval for numerical calculation was conducted. The result of BAAP is compared with the one from commercial 3D modeling software. Also the internal ballistic analysis was performed using the result of BAAP. In order to estimate the burning area and internal pressure with time, Chemical Equilibrium Analysis (CEA) was conducted with a composition of reduced smoke propellant. As a result, the web-averaged pressure was 5.34 MPa which is similar to the published research result.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.5 no.3
• /
• pp.41-51
• /
• 2001

Combustion characteristics of a KSR-III liquid rocket engine with split-triplet (F-O-O-F) type injector elements are investigated numerically from the viewpoints of engine performance and combustion flowfield. To evaluate numerical analysis of liquid rocket engine with radial type injector arrangement, 2-D axisymmetric and 3-D calculations are carried out and the prediction of engine performance for design and off-design conditions is in a good agreement with hot-firing tests. According to 2-D axisymmetric and 3-D calculations, the prediction error is 3∼5 % from the standpoint of performance. Numerical results of combustion characteristics calculated through 3-D analysis agree well with hot-firing tests qualitatively at injector plate. Decreasing impinging angle and changing radial type injector arrangement to H type injector arrangement reduce effectively local high-temperature region. Also, it is examined that those affect the performance seriously. In conclusion, it is revealed that both injector arrangement and impinging angle are critical parameters to affect the performance and combustion characteristics of the liquid rocket engine.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.11 no.5
• /
• pp.60-71
• /
• 2007

The startup characteristic of liquid propellant rocket engine should be focused on the stable ignition of combustion chamber and gas generator. Also, to lessen the propellants consumption during this period which doesn't contribute to the flight thrust, the engine has to be transferred to the nominal mode quickly. Because of the risk of test, it is impossible to develop all the startup cyclogram or the specifications of engine by test, so the precedent numerical approach is quite necessary. In this study we developed a mathematical model for the startup phenomena in a liquid rocket engine driven by gas generator-turbopump system based on the commercial 1-D flow system analysis program, Flowmaster. Using this program we proposed a methodology to obtain the specifications of turbine starter and the opening time of shutoff valves for the stable startup of the engine. To verify this methodology we qualitatively compared the analysis results to the typical startup curve of the published engine, then found it is quite well matched.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.11 no.5
• /
• pp.37-50
• /
• 2007

This study investigated the characteristics of spray generated by a liquid coaxial swirl injector used in a combustor of the liquid rocket engine. The linear stability analysis considered long and short wave was introduced in liquid sheet breakup. Through the hydrodynamic analysis the initial liquid sheet thickness spray angle and injection velocity were predicted. To evaluate the effect of turbulence model standard $k-{\varepsilon}$ and RNC $k-{\varepsilon}$ model were applied to numerical calculation and it was known that RNC $k-{\varepsilon}$ model was more applicable to predict spray characteristics. On the basis of this evaluation validation of the developed model was performed with swirl injector installed in LPRE and the predicted results of breakup length, spray angle, and SMD agreed well with experiments qualitatively and quantitatively.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.104-107
• /
• 2017

The power-pack combustion test of technology demonstration model(TDM0) for 9 tonf-class staged combustion cycle engine development was conducted in the Upper-stage Engine Test Facility(UETF) of Naro Space Center. The power-pack model of TDM0 was composed of a pre-burner, a turbo-pump and propellant supply systems without a main combustor. In the power-pack combustion test, we confirmed the linked working condition and verified the main functional variation of the power-pack for the engine system test.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.24 no.1
• /
• pp.17-23
• /
• 2020

Upper stage of launch vehicle mainly injects a lunar explorer from low earth orbit to the moon at a distance of 380,000 km. In foreign lunar explorer, the upper stage is separated from the explorer after the explorer is injected into the earth-moon transfer trajectory, and the lunar explorer then uses on-board propellant to carry out mid-course correction maneuvers and lunar orbit insertion maneuvers. This study describes a newly presented small liquid upper stage. Using a small liquid upper stage with a wet mass of 2.9 tonnes, the lunar explorer not only can be injected earth-moon transfer trajectory but also can be performed lunar orbit insertion. This study provides acceptable mass range of the lunar explorer and the scope of acceptable mission range also describes based on the launch from Naro Space Center.