• Aerospace Engineering and Technology
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• v.10 no.1
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• pp.89-97
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• 2011

This paper considers improved IGM (Iterative Guidance Mode), one of the explicit guidance algorithms, to determine the guidance algorithm for upper stages of a space launch vehicle. IGM, which has been employed successfully for the Saturn to put its payload into the parking orbit and lunar transfer orbit, is applied here for guidance of the launcher during the second and third stages. The orbit injection accuracy is evaluated through the 3-DOF computer simulations and an accurate prediction method, which can eliminate the prediction error of the downrange position at the orbit injection, is also proposed here.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.3
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• pp.97-103
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• 2012

The combustion chamber and nozzle of a liquid rocket engine should be protected from the high temperature combustion gas generated by the chamber. An upper-stage nozzle extension has a large expansion ratio, therefore, The light-weight refractory materials have been used since the weight impact on the launcher performance is crucial. Gas film cooling method was used before, but was not applicable nowadays. Ablative cooling method and radiative cooling method with niobium alloy, Ni-based superalloy and ceramic based composite have been used to this day.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.10
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• pp.853-861
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• 2012

This paper considers one of the explicit guidance algorithms, which has been proposed by Jaggers, to determine the closed-loop guidance algorithm for upper stages of a 3-staged space launch vehicle. Its commanded thrust vector is closer to the optimal solution when compared with that obtained by using the well-known Powered Explicit Guidance (PEG), which has been developed through the Space Shuttle program. Its performance is evaluated here by applying for guidance of the launcher during the second and third stages. Furthermore, to generate more precise guidance commands, it is attempted not to use the approximate formulas for the derivation of the original guidance law, and it is shown that performance is improved in comparison with the original.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.505-509
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• 2011

The combustion chamber and nozzle of a liquid rocket engine should be protected from the high temperature combustion gas generated by the chamber. An upper-stage nozzle extension has a large expansion ratio, therefore, The light-weight refractory materials have been used since the weight impact on the launcher performance is crucial. Gas film cooling and ablative cooling methods were used before, but were not applicable nowadays. Radiative cooling method with niobium alloy, Ni-based superalloy and ceramic based composite has been used to this day.

• Current Industrial and Technological Trends in Aerospace
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• v.9 no.1
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• pp.139-149
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• 2011

The combustion chamber and nozzle of a liquid rocket engine need thermal protection against the high temperature combustion gas. The nozzle extension of a high-altitude engine also has to be compatible with high temperature environment and several kinds of cooling methods including gas film cooling, ablative cooling and radiative cooling are used. Especially for an upper-stage nozzle extension having a large expansion ratio, the weight impact on the launcher performance is crucial and it necessitated the development of light-weight refractory material. The present survey on the nozzle extension materials employed in the liquid rocket engines of USA, Russia and European Union has revealed a trend that the heavier metals like stainless steels and titanium alloys are being substituted with light weight carbon fiber or ceramic matrix composite materials.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.50 no.12
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• pp.867-875
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• 2022

The preliminary mission analysis of a lunar lander, which is mounted on the upper stage of a Korean space launch vehicle, is performed when landing on the moon through a trans-lunar injection maneuver after being injected into the earth's low orbit by th launcher in this paper. Both direct landing and orbital landing methods, which have each advantage and disadvantages, are applied and their transfer orbit characteristics are analyzed according to the launch date when launching in lunar October 2030. We also analyzed the launch dates which satisfying eclipse conditions, solar elevation conditions, and tracking time intervals such as the US lunar lander Surveyor-1. The obtained results show that the most appropriate launch date is the 4th day of lunar October in case of direct landing method, and the 3rd day in case of indirect landing method, since the argument of perigee of the trans-lunar injection orbit and eclipse conditions are favorable in the dates.

• Nuclear Engineering and Technology
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• v.55 no.8
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• pp.2941-2951
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• 2023

The optimal minimum ECCD power is evaluated numerically for completely suppressing the 3/2 and 2/1 NTMs in the CFETR hybrid scenario. For two typical frequencies of ECCD sources launching from two upper launcher (UL) ports, f_ec = 210 GHz and 240 GHz with O1-mode, UL₁: (R_i, Z_i) = (8.47, 5.7) m and UL₂: (R_i, Z_i) = (8.2, 4.5) m, higher frequency of ECCD source launching from the UL₂ port is better than that low frequency counterpart from the UL₁ port. Using 240 GHz ECCD source launching from the UL₂ port, the minimum power required to fully suppress the two NTMs with precise ECCD alignment is 12.4 MW and 16.7 MW, respectively. When good alignment cannot be achieved, the results suggest that the misalignment should not exceed 0.02α, preferably 0.015α, corresponding to 4.4 cm and 3.3 cm. Considering engineering difficulty of high-frequency gyrotron sources, the optimal minimum ECCD power with the 210 GHz source launching from the UL₂ port is 17.9 MW and 20.6 MW for completely suppressing the 3/2 and 2/1 NTMs, respectively. In view of this, it is a good choice to select the 210 GHz ECCD source launching from the UL₂ port in the short and medium term.

• Journal of Astronomy and Space Sciences
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• v.23 no.4
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• pp.355-372
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• 2006

Mission opportunities and trajectory characteristics for the future Korean Mars mission have designed and analyzed using KSIV-III(Korea Space Launch Vehicle-III). Korea's first space center, 'NARO space center' is selected as a launch site. For launch opportunities, year 2033 is investigated under considering the date of space center's completion with KSLV series development status. Optimal magnitude of various maneuvers, Trans Mars Injection (TMI) maneuver, Trajectory Correction Maneuver (TCM), Mars Orbit Insertion (MOI) maneuver and Orbit Trim Maneuver(OTM), which are required during the every Mars mission phases are computed with the formulation of nonlinear optimization problems using NPSOL software. Finally, mass budgets for upper stage (launcher for KSIV-III and spacecraft are derived using various optimized maneuver magnitudes. For results, daily launch window from NARO space center for successful Korean Mars mission is avaliable for next 27 minutes starting from Apr. 16. 2033. 12:17:26 (UTC). Maximum spacecraft gross mass which can delivered to Mars is about 206kg, with propellant mass of 109kg and structure mass of 97kg, when on board spacecraft thruster's Isp is assumed to have 290 sec. For upper stage, having structure ratio of 0.15 and Isp value of 280 sec, gross mass is about 1293kg with propellant mass of 1099kg and structure mass of 194kg. However, including 10% margins to computed optimal maneuver values, spacecraft gross mass is reduced to about 148kg with upper stage's mass of 1352kg. This work will give various insights, requiring performances to developing of KSIV-III and spacecraft design for future Korean Mars missions.

• Journal of the Korea Society for Simulation
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• v.21 no.1
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• pp.81-88
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• 2012

In the M&S filed, The Battle Lab is available for acquisition, design, development tool, validation test, and training in the weapon system of development process. Recently, the Battle Lab in the military of Korea is still in an early stage, in spite of importance of battle lab construction. In the environment of network centric warfare, a practical use of the M&S which is connecting live, virtual and constructive model can be applied to all field of System Engineering process. It is necessary thar the Battle Lab is not restricted by time and space, and is possible for the technical implementation. In this paper, to guarantee the interoperability of live and virtual simulation, virtual simulators connect live simulators by using the tactical data link. To guarantee the interoperability of virtual and constructive simulation, both virtual simulators and constructive simulators use the RTI which is the standard tool of M&S. We propose the System that constructed the Air Defence Battle Lab. In case of the approach of target tracks, The Air Defence Battle Lab is the system for the engagement based on a command of an upper system in the engagement weapon system. Constructive simulators which are target track, missile, radar, and launcher simulator connect virtual simulators which are MCRC, battalion, and fire control center simulators using the RPR-FOM 1.0 that is a kind of RTI FOM. The interoperability of virtual simulators and live simulators can be guaranteed by the connection of the tactical data links which are Link-11B and ATDL-1.