• Aerospace Engineering and Technology
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• v.7 no.1
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• pp.91-98
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• 2008

Since the 1st space race between the United States and Soviet Union during the 1960s, we are competing 2nd space race to occupy the Lunar territory. Since the United States announced to construct the Lunar Base by the end of 2020, ED, Japan, and China launched Lunar explorers successfully. Even India is planning to launch a Lunar explorer in 2008. Korean government also announced that the Korea will launch first Lunar explorer in 2020. In this research Lunar mission trajectory design which will be fundamental data for Lunar mission with variable low thrust and Lunar mission trajectory which has a similar mission specification to SMART-1 are presented.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.4
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• pp.357-367
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• 2008

Preparing for future Korean Lunar missions, preliminary Lunar mission design software is developed using a impulsive thrusting method. Developed software is capable of design and analysis every required mission phases to design Lunar mission, including the Earth departure, Lunar transfer, Lunar arrival and mission operation phases. Also, assuming that KSLV-II is selected as a launch vehicle, future Korean Lunar explorer's mass budget is estimated based on driven optimal trajectory characteristics. Tracking analysis is also performed using Deep Space Network including angle geometry analysis between Earth - Moon - Lunar explorer - Sun which are very important for communication, solar panel pointing strategy and eclipse analysis when Lunar missions are under designing phase.

• Aerospace Engineering and Technology
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• v.8 no.1
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• pp.108-116
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• 2009

This paper considers the trajectory design problem for a lunar orbiter when launched by KSLV-II. KSLV-II puts its kick motor stage and lunar orbiter into a low earth orbit, and then the kick motor stage performed the translunar injection. To simulate more realistic situations, TLI (Trans-Lunar Injection) and LOI (Lunar Orbit Injection) maneuvers are modeled as finite burns. The feasibility of the lunar mission by KSLV-II are confirmed by the numerical results that show the reasonable required-velocity and propellant usage.

• Current Industrial and Technological Trends in Aerospace
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• v.8 no.1
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• pp.62-76
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• 2010

In this paper, recent researches about the unmanned lunar lander development are studied. Objectives of this study are to derive research trends and to identify key design activities especially in early design phase of unmanned lunar lander. Case study covers SELENE-2 of Japan, LEDA and MoonNEXT of ESA, and small and modular spacecraft approach of NASA. Lunar lander concepts proposed for the International Lunar Network Anchor Nodes are also studied. For each lunar lander program, mission requirements are summarized and mission design results are reviewed. Approaches of safe lunar landing including design of navigation, guidance and control, combination of sensors, derived sensor and propulsion performance requirements are also analyzed.

• Journal of Astronomy and Space Sciences
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• v.32 no.3
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• pp.257-268
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• 2015

Characteristics of delta-V requirements for deploying an impactor from a mother-ship at different orbital altitudes are analyzed in order to prepare for a future lunar CubeSat impactor mission. A mother-ship is assumed to be orbiting the moon with a circular orbit at a 90 deg inclination and having 50, 100, 150, 200 km altitudes. Critical design parameters that are directly related to the success of the impactor mission are also analyzed including deploy directions, CubeSat flight time, impact velocity, and associated impact angles. Based on derived delta-V requirements, required thruster burn time and fuel mass are analyzed by adapting four different miniaturized commercial onboard thrusters currently developed for CubeSat applications. As a result, CubeSat impact trajectories as well as thruster burn characteristics deployed at different orbital altitudes are found to satisfy the mission objectives. It is concluded that thrust burn time should considered as the more critical design parameter than the required fuel mass when deducing the onboard propulsion system requirements. Results provided through this work will be helpful in further detailed system definition and design activities for future lunar missions with a CubeSat-based payload.

• Journal of Astronomy and Space Sciences
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• v.28 no.3
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• pp.203-216
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• 2011

To prepare for a future Korean lunar orbiter mission, semi-optimal lunar capture orbits using finite thrust are designed and analyzed. Finite burn delta-V losses during lunar capture sequence are also analyzed by comparing those with values derived with impulsive thrusts in previous research. To design a hypothetical lunar capture sequence, two different intermediate capture orbits having orbital periods of about 12 hours and 3.5 hours are assumed, and final mission operation orbit around the Moon is assumed to be 100 km altitude with 90 degree of inclination. For the performance of the on-board thruster, three different performances (150 N with $I_{sp}$ of 200 seconds, 300 N with $I_{sp}$ of 250 seconds, 450 N with $I_{sp}$ of 300 seconds) are assumed, to provide a broad range of estimates of delta-V losses. As expected, it is found that the finite burn-arc sweeps almost symmetric orbital portions with respect to the perilune vector to minimize the delta-Vs required to achieve the final orbit. In addition, a difference of up to about 2% delta-V can occur during the lunar capture sequences with the use of assumed engine configurations, compared to scenarios with impulsive thrust. However, these delta-V losses will differ for every assumed lunar explorer's on-board thrust capability. Therefore, at the early stage of mission planning, careful consideration must be made while estimating mission budgets, particularly if the preliminary mission studies were assumed using impulsive thrust. The results provided in this paper are expected to lead to further progress in the design field of Korea's lunar orbiter mission, particularly the lunar capture sequences using finite thrust.

• 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.

• Journal of Astronomy and Space Sciences
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• v.27 no.3
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• pp.195-204
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• 2010

The optimal Earth-Moon transfer trajectory considering spacecraft's visibility from the Daejeon ground station visibility at both the trans lunar injection (TLI) and lunar orbit insertion (LOI) maneuvers is designed. Both the TLI and LOI maneuvers are assumed to be impulsive thrust. As the successful execution of the TLI and LOI maneuvers are crucial factors among the various lunar mission parameters, it is necessary to design an optimal lunar transfer trajectory which guarantees the visibility from a specified ground station while executing these maneuvers. The optimal Earth-Moon transfer trajectory is simulated by modifying the Korean Lunar Mission Design Software using Impulsive high Thrust Engine (KLMDS-ITE) which is developed in previous studies. Four different mission scenarios are established and simulated to analyze the effects of the spacecraft's visibility considerations at the TLI and LOI maneuvers. As a result, it is found that the optimal Earth-Moon transfer trajectory, guaranteeing the spacecraft's visibility from Daejeon ground station at both the TLI and LOI maneuvers, can be designed with slight changes in total amount of delta-Vs. About 1% difference is observed with the optimal trajectory when none of the visibility condition is guaranteed, and about 0.04% with the visibility condition is only guaranteed at the time of TLI maneuver. The spacecraft's mass which can delivered to the Moon, when both visibility conditions are secured is shown to be about 534 kg with assumptions of KSLV-2's on-orbit mass about 2.6 tons. To minimize total mission delta-Vs, it is strongly recommended that visibility conditions at both the TLI and LOI maneuvers should be simultaneously implemented to the trajectory optimization algorithm.

• Journal of Astronomy and Space Sciences
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• v.41 no.1
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• pp.43-60
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• 2024

Korea Pathfinder Lunar Orbiter (KPLO) is South Korea's first space exploration mission, developed by the Korea Aerospace Research Institute. It aims to develop technologies for lunar exploration, explore lunar science, and test new technologies. KPLO was launched on August 5, 2022, by a Falcon-9 launch vehicle from cape canaveral space force station (CCSFS) in the United States and placed on a ballistic lunar transfer (BLT) trajectory. A total of four trajectory correction maneuvers were performed during the approximately 4.5-month trans-lunar cruise phase to reach the Moon. Starting with the first lunar orbit insertion (LOI) maneuver on December 16, the spacecraft performed a total of three maneuvers before arriving at the lunar mission orbit, at an altitude of 100 kilometers, on December 27, 2022. After entering lunar orbit, the commissioning phase validated the operation of the mission mode, in which the payload is oriented toward the center of the Moon. After completing about one month of commissioning, normal mission operations began, and each payload successfully performed its planned mission. All of the spacecraft operations that KPLO performs from launch to normal operations were designed through the system operations design process. This includes operations that are automatically initiated post-separation from the launch vehicle, as well as those in lunar transfer orbit and lunar mission orbit. Key operational procedures such as the spacecraft's initial checkout, trajectory correction maneuvers, LOI, and commissioning were developed during the early operation preparation phase. These procedures were executed effectively during both the early and normal operation phases. The successful execution of these operations confirms the robust verification of the system operation.

• 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.