• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.92-97
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• 2012

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility (CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The oxidizer supply system modeling using AMESim was performed based on the results of the detailed design, and the oxidizer supply characteristics was analyzed in this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.56-61
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• 2012

Fuel-side cold-flow tests were performed on the technology demonstration model of a 75 ton-class liquid rocket engine combustion chamber for the first stage of the Korea space launch vehicle II. Pressure drop in the cooling channels of the combustion chamber was measured by changing fuel mass flow rate through a pressure regulating system. Pressure drop in each segment of the chamber could be obtained and a lot of pressure drop was caused by high flow velocity in the nozzle throat segment. The accuracy of a hydraulic analysis method for calculating a pressure loss in cooling channels could be verified by applying it to the cold-flow test conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.4
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• pp.89-95
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• 2017

The improvement of a flow coefficient for the recirculation chill-down flow in a main oxidizer shut-off valve has been presented. The flow coefficient, which is mainly affected by the recirculation outlet port size and the configuration inside the valve, has been predicted with measured flow coefficient values. The comparison of experimentally measured flow coefficient with the predicted value shows the effect of valve inside configuration on the flow coefficient. Consequently, the flow coefficient is twice the previous value and about 75% of the pressure loss assigned to the main oxidizer shut-off valve can be used for additional pressure losses for other components in the recirculation chill-down system of a launch vehicle.

• Fire Science and Engineering
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• v.33 no.5
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• pp.13-18
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• 2019

In this study, the soot formation characteristics of Jet-A1 liquid fuel droplet flames were investigated by measuring the soot concentration under atmospheric conditions similar to the working environment of the Korea Space Launch Vehicle (KSLV) To obtain the desired atmospheric conditions, the oxygen concentration in the combustion chamber was maintained at 30% and the pressure was varied between 0.1 and 0.06 MPa. The full-field light extinction technique was used to measure the concentration of soot particles generated by applying the identical to 2-mm-diameter Jet-A1 fuel droplets. The soot concentration of the Jet-A1 droplet flames was the highest in the nitrogen-substituted atmosphere and the lowest in the carbon dioxide-substituted atmosphere, despite the pressure. the pressure was decreased the measured soot concentrations reduced as a function of Pⁿ.

• Bulletin of the Korean Space Science Society
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• 2004.10b
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• pp.372-375
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• 2004

The UV radiometer payload was launched successfully from the west coastal area of Korea Peninsula aboard KSR-III on 28, Nov 2002. KSR-III was the Korean third generation sounding rocket and was developed as intermediate step to larger space launch vehicle with liquid propulsion engine system. UV radiometer onboard KSR-III consists of UV and visible band optical phototubes to measure the direct solar attenuation during rocket ascending phase. For UV detection, 4 channel of sensors were installed in electronics payload section and each channel has 255, 290, 310nm center wavelengths, respectively. 450nm channel was used as reference for correction of the rocket attitude during the flight. Transmission characteristics of all channels were calibrated precisely prior to the flight test at the Optical Lab. in KARI (Korea Aerospace Research Institute). During a total of 231s flight time, the onboard data telemetered to the ground station in real time. The ozone column density was calculated by this telemetry raw data. From the calculated column density, the vertical ozone profile over Korea Peninsula was obtained with sensor calibration data. Our results had reasonable agreements compared with various observations such as ground Umkhr measurement at Yonsei site, ozonesonde at Pohang site, and satellite measurements of HALOE and POAM. The sensitivity analysis of retrieval algorithm for parameters was performed and it was provided that significant error sources of the retrieval algorithm.

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.149-158
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• 2006

A review has been made on what kind of method can be applied to predict the dynamic derivatives of the separated PLF(Payload Fairing) halves of a launch vehicle in consideration of technology and budget. An optimal approach is selected considering the geometric characteristics of the PLF halves, the aerodynamic conditions and the required accuracy. The time history of aerodynamic force/moment coefficients are obtained for the forced harmonic motions by solving the unsteady Euler equations derived with respect to the inertial reference frame. and the dynamic derivatives are deduced by integration of the aerodynamic coefficients for one period. In this research, the dynamic derivatives are presented for 0.6$\leq$ M $\leq$2.0, $-180^{\circ}$ $\leq$$\alpha$ $\leq$$180^{\circ}$ and $-90 ^{\circ}$$\leq$$\beta$$\leq$$90 ^{\circ}$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.585-589
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• 2010

It is important for the liquid rocket engine to meet the exact performance requirements in order to guarantee the successful mission of the launch vehicle. Usually, a ground combustion test for the engine is conducted to reduce the performance error and for the tuning. For the gas-generator (GG) cycle engine, this adjustment process can be easily tuned by means of the control valves. A linearized correlation between the process parameters of the control - the combustion chamber pressure and the mixture ratio of engine - and the independent parameter of the control- rotational angle of the control valve - could be suitable to reduce the tuning errors. Also this linearity can reduce the effort for the tuning and make the process more explicit by ensuring a more intuitive control. In this point, we proposed an algorithm in the frame of the in-house-developed program to obtain the control valves' inherent characteristics which satisfy the linearity.

• Aerospace Engineering and Technology
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• v.1 no.2
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• pp.57-65
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• 2002

The Electrical Power System (EPS) shall supply required power to maintain spacecraft and payload during the mission. The EPS sizing are based on space environment, satellite mission and lifetime, and allocated budgets. The type of the primary and secondary power is determined according to satellite design-level and allocated subsystem budgets. The design of EPS has closely related to system and others' subsystems design. To supply the sufficient power to the satellite, the implementation of the larger power source and energy storage is impossible actually. And there will be some problems of the attitude control of the satellite, the handling power capability of the electronic boxes, and launch vehicle selection caused by EPS oversizing. Also, the thermal control is not easy in the space by extra power. And the maintenance of the satellite within the specific orbit from orbit-drag is a big design burden of the thruster. So the various technologies have been developed to optimize the EPS sizing and to operate the power system efficiently.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.165-172
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• 2004

LE-7A is the main engine of the H-IIA launch vehicle. Under its development, the nozzle suffered from two troubles during startup and shutdown transients of the engine. One is a large side load, which damages the actuator of the nozzle, and the other is damage on regenerative cooling tubes due to high heat load. It has been considered that these problems are caused by a peculiar separation pattern called Restricted Shock Separation (RSS). RSS is observed in several rocket nozzles, for example, LE-7A nozzle, Vulcain nozzle and so on. Their contours are not conventional truncated perfect (TP) nozzle - LE-7A nozzle is a compressed truncated perfect (CTP) nozzle and Vulcain nozzle is a thrust optimized (TO) nozzle. Although it is believed that the occurrence of RSS is affected by the nozzle contour, the mechanisms are not clarified sufficiently yet. In the present paper, a parametric numerical study is carried out to investigate the mechanisms of the occurrence of RSS in CTP nozzles during startup transient. The results show that RSS is caused by the adverse pressure gradient downstream of the Mach disk. The adverse pressure gradient is caused by the interaction of the pressure wave and Mach disk. The method to avoid the occurrence of RSS is also examined. A small step inside the nozzle affects the position of the separation point and prevents RSS. The result shows that the possibility that RSS can be suppressed by controlling the position of the separation point.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.6
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• pp.113-119
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• 2013

A main oxidizer shut-off valve (MOV) controls the supply of cryogenic liquid oxygen to the combustion chamber of liquid rocket engines by on/off operations. The main subjects to be introduced are not only the valve transient response during valve on/off procedures but also the characteristics of pneumatic and seat/poppet parts as core technologies in the development of the MOV, which is expected to be adopted for the Korea Space Launch Vehicle II. It is shown that the analytical prediction of the transient valve travel is in good agreement with experimental results. Friction and elastic forces on the valve moving part are quantitatively evaluated by structural analysis.