• International Journal of Aeronautical and Space Sciences
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• v.13 no.4
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• pp.474-483
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• 2012

In this paper, a collision avoidance maneuver was sought for low Earth orbit (LEO) and geostationary Earth orbit (GEO) satellites maintained in a keeping area. A genetic algorithm was used to obtain both the maneuver start time and the delta-V to reduce the probability of collision with uncontrolled space objects or debris. Numerical simulations demonstrated the feasibility of the proposed algorithm for both LEO satellites and GEO satellites.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.9
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• pp.60-66
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• 2006

Now, on developing COMS(Communication, Ocean and Meteorological Satellite) has solar panel on the south panel only. Therefore, the wheel off-loading has to be performed periodically to reduce a induced momentum energy by a asymmetric solar panel. One of two East/West station keeping maneuver to correct simultaneously longitude and eccentricity, orbit corrections may be performed during one of the two wheel off-loading manoeuvres per day to get enough observation time for meteorological and ocean sensor. In this paper, we applied a linearized orbit maneuver equation to acquire maneuver time and delta-V. Nonlinear simulation for the station keeping is performed and compared with general station keeping strategy for fuel reduction.

• Journal of Astronomy and Space Sciences
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• v.38 no.1
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• pp.65-82
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• 2021

For the vast majority of geostationary satellites currently in orbit, station keeping activities including orbit determination and maneuver planning and execution are ground-directed and dependent on the availability of ground-based satellite control personnel and facilities. However, a requirement linked to satellite autonomy and survivability in cases of interrupted ground support is often one of the stipulated provisions on the satellite platform design. It is especially important for a geostationary military-purposed satellite to remain within its designated orbital window, in order to provide reliable uninterrupted telecommunications services, in the absence of ground-based resources due to warfare or other disasters. In this paper we investigate factors affecting the robustness of a geostationary satellite's orbit in terms of the maximum duration the satellite's station keeping window can be maintained without ground intervention. By comparing simulations of orbit evolution, given different initial conditions and operations strategies, a variation of parameters study has been performed and we have analyzed which factors the duration is most sensitive to. This also provides valuable insights into which factors may be worth controlling by a military or civilian geostationary satellite operator. Our simulations show that the most beneficial factor for maximizing the time a satellite will remain in the station keeping window is the operational practice of pre-emptively loading East-West station keeping maneuvers for automatic execution on board the satellite should ground control capability be lost. The second most beneficial factor is using short station keeping maneuver cycle durations.

• 제어로봇시스템학회:학술대회논문집
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• 2000.10a
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• pp.458-458
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• 2000

Quick evaluations of two in-plane orbit maneuvers using small see of real-time CPS navigation solutions were peformed lot the KOMPSAT-1 spacecraft operation. Real-time GPS navigation solutions of the KOMPSAT-1 were collected during the Korean Ground Station(KGS) pass. Only a few sets of position and velocity data after completion of the thruster firing were used for the quick maneuver evaluations. The results were used for antenna pointing data predictions for the next station contact. Normal orbit maneuver evaluations using large see of playback GPS navigation solutions were also performed and the result were compared with the quick evaluation results.

• Journal of Astronomy and Space Sciences
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• v.14 no.2
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• pp.355-366
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• 1997

This paper presents a comprehensive analytical approach for determining key maneuver parameters associated with the acquisition and maintenance of the ground track for a low-earth orbit. A livearized model relating changes in the drift rate of the ground track directly to changes in the orbital semi-major axis is also developed. The effect of terrestrial atmospheric drag on the semi-major axis is also explored, being quantified through an analytical expression for the decay rate as a function of density. The non-singular Lagrange planetary equations, further simplified for nearly circular orbits, provide the desired relationships between the corrective in-plane impulsive velocity increments and the corresponding effects on the orbit elements. The resulting solution strategy offers excellent insight into the dynamics affecting the timing, magnitude, and frequency of these maneuvers. Simulations are executed for the ground track acquisition and maintenance maneuver as a pre-flight planning and analysis.

• Journal of Astronomy and Space Sciences
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• v.12 no.2
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• pp.265-274
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• 1995

The purpose of perigee kick motor firing is to place a satellite into transfer orbit and that of apogee kick motor firing is to place the satellite into geosynchonous orbit in order to increase the semi-major axis of the transfer orbit and reduce the inclination of the transfer orbit. Because apogee motor firing is always accompanied with injection errors, the satellite is not placed into geosynchonous orbit but into a near-geosynchonous orbit, also knows as a drift orbit. Thus, the orbital maneuver to correct drift orbit into gteosynchonous orbit is required, this maneuver is called the station acquisition. For reduction of expenditure and performance of mission, we estimate $\Delta$V budget and required fuel allowance for station acquisition. As the uncertainty of drift orbit by injection error of perigee and apogee kick motor firing prevents us from obtaining exact $\Delta$V budget, statistical Monte Carlo simulation technique is used in order to get optimal $\Delta$V budget and required fuel allowance with a probability of 99%. With respect to Korea satellite launched by Delta-2 launch vehicle in 1995, Monte Carlo analysis is used in order to get various orbital parameters, $\Delta$V budget and required fuel allowance for station acquisition with a probability of 99%.

• Journal of Astronomy and Space Sciences
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• v.34 no.4
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• pp.331-342
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• 2017

To ensure the successful launch of the Korea pathfinder lunar orbiter (KPLO) mission, the Korea Aerospace Research Institute (KARI) is now performing extensive trajectory design and analysis studies. From the trajectory design perspective, it is crucial to prepare contingency trajectory options for the failure of the first lunar brake or the failure of the first lunar orbit insertion (LOI) maneuver. As part of the early phase trajectory design and analysis activities, the required time of flight (TOF) and associated delta-V magnitudes for each recovery maneuver (RM) to recover the KPLO mission trajectory are analyzed. There are two typical trajectory recovery options, direct recovery and low energy recovery. The current work is focused on the direct recovery option. Results indicate that a quicker execution of the first RM after the failure of the first LOI plays a significant role in saving the magnitudes of the RMs. Under the conditions of the extremely tight delta-V budget that is currently allocated for the KPLO mission, it is found that the recovery of the KPLO without altering the originally planned mission orbit (a 100 km circular orbit) cannot be achieved via direct recovery options. However, feasible recovery options are suggested within the boundaries of the currently planned delta-V budget. By changing the shape and orientation of the recovered final mission orbit, it is expected that the KPLO mission may partially pursue its scientific mission after successful recovery, though it will be limited.

• Aerospace Engineering and Technology
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• v.13 no.1
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• pp.96-107
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• 2014

Korea Aerospace Research Institute has a plan to launch experimental lunar orbiter in 2017, and lunar orbiter and lander in 2020. In the mission planning phase, LOI(Lunar Orbit Insertion) maneuver strategy should be designed using finite burn model because on-board propulsion system of lunar orbiter in finite burn type. LOI maneuver plan and amount of required ${\Delta}V$ using finite burn model depend on the spacecraft attitude at burn, a type of propellant, thrust level and burn timing. This paper describes the LOI maneuver of lunar orbiter of foreign space agency and then comes up with the LOI maneuver plan of Korean lunar orbiter. Adequate thrust level and burn duration of Korean lunar orbiter also present by performing simulation.

• ETRI Journal
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• v.30 no.6
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• pp.774-782
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• 2008

An operational orbit determination (OD) and prediction system for the geostationary Communication, Ocean, and Meteorological Satellite (COMS) mission requires accurate satellite positioning knowledge to accomplish image navigation registration on the ground. Ranging and tracking data from a single ground station is used for COMS OD in normal operation. However, the orbital longitude of the COMS is so close to that of satellite tracking sites that geometric singularity affects observability. A method to solve the azimuth bias of a single station in singularity is to periodically apply an estimated azimuth bias using the ranging and tracking data of two stations. Velocity increments of a wheel off-loading maneuver which is performed twice a day are fixed by planned values without considering maneuver efficiency during OD. Using only single-station data with the correction of the azimuth bias, OD can achieve three-sigma position accuracy on the order of 1.5 km root-sum-square.

• Journal of Astronomy and Space Sciences
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• v.34 no.2
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• pp.127-138
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• 2017

The current study designs the mission orbit of the lunar CubeSat spacecraft to measure the lunar local magnetic anomaly. To perform this mission, the CubeSat will impact the lunar surface over the Reiner Gamma swirl on the Moon. Orbit analyses are conducted comprising ${\Delta}V$ and error propagation analysis for the CubeSat mission orbit. First, three possible orbit scenarios are presented in terms of the CubeSat's impacting trajectories. For each scenario, it is important to achieve mission objectives with a minimum ${\Delta}V$ since the CubeSat is limited in size and cost. Therefore, the ${\Delta}V$ needed for the CubeSat to maneuver from the initial orbit toward the impacting trajectory is analyzed for each orbit scenario. In addition, error propagation analysis is performed for each scenario to evaluate how initial errors, such as position error, velocity error, and maneuver error, that occur when the CubeSat is separated from the lunar orbiter, eventually affect the final impact position. As a result, the current study adopts a CubeSat release from the circular orbit at 100 km altitude and an impact slope of $15^{\circ}$, among the possible impacting scenarios. For this scenario, the required ${\Delta}V$ is calculated as the result of the ${\Delta}V$ analysis. It can be used to practically make an estimate of this specific mission's fuel budget. In addition, the current study suggests error constraints for ${\Delta}V$ for the mission.