Publication

We present a system-level design for a dual near-Earth asteroid mission that combines reconnaissance and kinetic impact within a single planetary-defense architecture. The concept stages a chemically propelled spacecraft in the Sun–Earth L1 region and targets two asteroids with consecutive close Earth approaches: 2024 YR4, which approaches from the sunward direction and is difficult to characterize from the ground, and 2021 PC7, which offers a favorable geometry for a controlled kinetic impact. A multi-impulse trajectory, optimized in the Sun–Earth circular restricted three-body problem using a multiple-point differential correction algorithm, enables a ballistic transfer and reconnaissance flyby of 2024 YR4 near SEL1, followed by a nine-month transfer to a terminal impact on 2021 PC7. We develop a preliminary 300-kg-class spacecraft configuration sized to implement this trajectory, including orbit and attitude maneuvering subsystems, power and communications, and a dedicated payload suite featuring a rotating-mirror imaging system and spectrometer for high-speed flyby observations. The spacecraft also carries a 6U CubeSat, deployed shortly before the 2021 PC7 encounter, to perform a close flyby of the impact site and image the flash, ejecta plume, and crater formation. The paper describes the overall mission architecture, reference trajectory, spacecraft system design, and detailed operational concepts for the 2024 YR4 flyby, the 2021 PC7 impact, and coordinated use of the primary and CubeSat payloads. This study illustrates how SEL1 staging, dual-target operations, and small secondary spacecraft can be integrated into a feasible mission that exercises key elements of an end-to-end planetary-defense response.

This manuscript presents the concept of operations and trajectory architecture for a single-spacecraft, dual-target, Near-Earth-Object (NEO) planetary-defense mission: a close reconnaissance flyby of 2024 YR4, followed by a kinetic impact on 2021 PC7 during their close Earth approach. The baseline mission architecture employs a Sun–Earth L1 ballistic transfer timed to meet 2024 YR4 near the Sun-Earth L1 region in December of 2032. This flyby location reduces the spacecraft–asteroid relative velocity versus an Earth-vicinity flyby and leverages launch-vehicle hyperbolic excess energy (C3), thereby lowering the DV required for the subsequent 2021 PC7 impact leg. The Earth-to-2024 YR4 transfer is constructed by following the stable manifold of a Sun–Earth L1 periodic orbit that intersects 2024 YR4’s trajectory at the desired flyby epoch and location for flyby. Post-flyby maneuvers for the 2021 PC7 kinetic impact are solved as an N-impulse optimization problem designed to minimize the total $\Delta V$ subject to the encounter window. We present the end-to-end design flow, key trades, and the Sun-Asteroid-Space angle, demonstrating the feasibility of a low-DV planetary defense mission.

Space Situational Awareness (SSA) capabilities are becoming increasingly critical as space becomes more accessible and activities intensify. SSA is essential for operations, planning, and strategic-political decision support globally. In parallel with worldwide efforts by various organizations, the Korea Advanced Institute of Science and Technology (KAIST) and the Korea Astronomy and Space Science Institute (KASI) have developed risk analysis tools, with a current focus on 3D cell-based analysis software. This paper introduces the baseline model and implementation details of the developed software. We also present the results and implications of numerical studies focusing on low Earth orbit. These studies analyze the impact of input parameters and investigate the evolution of the space collision environment, providing insights for parameter selection and demonstrating the applicability of the developed software.

Human missions to the lunar surface require transporting various commodities from Earth to the Moon. Commodities arriving on the Moon before humans can support the operations effectively. Therefore, selecting orbital sequences with appropriate propulsion systems (e.g., chemical or electric) and associated trajectories is essential for the success of human lunar missions. Each commodity must be optimized considering the limited time of flight (TOF) and the cargo mass, to utilize allowable launch windows before manned missions efficiently, ultimately reducing the number of launches. This paper addresses an optimization framework for a multicommodity lunar campaign with resonance orbits and hybrid propulsion. The Pareto front of TOF and fuel consumption is calculated by solving multiple mixed-integer linear programming models from precalculated trajectory optimization results. The optimal logistics planning, addressing multiple commodities using the calculated Pareto front, is formulated as a mixed integer programming. A realistic multicommodity lunar exploration mission utilizing hybrid propulsion and multiple resonance orbits is used as a case study to validate the proposed methodology.

This study focuses on optimizing low-thrust trajectories for a spacecraft to achieve an inclined Sun-Earth L4 periodic orbit. The optimization is formulated as an indirect optimization problem, based on the Euler-Lagrange equations of motion. The objective is to minimize the total propellant mass, following the Pontryagin Minimum Principle, which maximizes the spacecraft's mass upon arrival at the Sun-Earth L4 point. The analysis includes two key optimal control problems: Sun-Earth L4 insertion and inclination-pumping optimal control problem. The Sun-Earth L4 insertion optimal control problem solves the optimal thrusting direction and throttling required to stop at the Sun-Earth L4 with the desired ecliptic inclination after launch. The inclination-pumping optimal control problem solves the optimal thrusting direction and throttling required to move the spacecraft from a low to a high-inclination orbit about Sun-Earth L4. The first continuation strategy is transitioning the spacecraft trajectory from energy-optimal to fuel-optimal solutions. Then, a second continuation strategy is employed to decrease and increase the maximum thrust level, which generates the control surfaces that reveal the relationship between fuel-optimal trajectories and thrust levels. The test cases involve a 1,500 kg spacecraft equipped with a 200 mN electric thruster powered by solar arrays that provide 3 kW end‑of‑life—2-kW of which is required, leaving a steady 1 kW margin. These cases analyze the mass at arrival for various initial inclinations and maneuver sequences. The analysis performed in the case study section targets 14.5 degrees inclined Sun-Earth L4 periodic orbit with several intermediate inclinations. Optimal launch windows for high-latitude solar surface observations are calculated for each trajectory type, accounting for the tilt angle of the Sun's rotational axis from the ecliptic frame.

We introduce the long-term Earth magnetosphere science orbits designed to maintain a fixed orientation relative to Earth’s magnetosphere over extended durations. By leveraging the Earth-Moon resonant orbits, the spacecraft’s argument of periapsis is aligned with the orientation of Earth’s magnetosphere, thereby enabling continuous observations. Three specific Earth–Moon resonant orbits, characterized by distinct values of the Jacobi integral, are identified to exhibit these properties of stable, magnetosphere-aligned evolution. This approach facilitates sustained monitoring of large-scale magnetospheric dynamics and opens new opportunities for focused science objectives. These include studying the interaction between the Earth and the Moon in shaping magnetospheric boundaries and probing magnetospheric vortices and other transient phenomena. The resultant long-term vantage point—achieved through careful resonance and orbital design—offers a platform for future space weather research, multi-point observations of Earth-Moon system interactions, and advancing magnetospheric science missions overall.

With the recent surge in the number of objects in orbit, the collision risk to space assets has grown rapidly. Accurately assessing this risk to support safe operations has therefore become increasingly important. We develop a 3D cell model—a space environment modeling technique for collision assessment—and validate it by comparing our spatial density calculations in low Earth orbit and collision risk results for KITSAT against those produced by MASTER, the 3D grid-cell model developed by ESA.

The multiple Mars gravity-assist trajectory is compared to the phasing trajectory for placing a spacecraft in a circular Sun-Earth L4 orbit with a 1 AU semi-major axis and inclinations of 10 and 14.5 degrees relative to the ecliptic plane. The gravity-assist maneuvers are treated as instantaneous velocity changes using a zero-sphere-of-influence model. The trajectory is optimized for two potential launch vehicles (Falcon 9 and Falcon Heavy) to achieve the desired orbit with minimal C3 energy. Through trajectory analysis based on various launch vehicles and their C3-based payload capacities, it was found that the multiple Mars gravity-assist trajectories are outperformed by the phasing trajectory at a10 degrees inclination but are additions to the Pareto optimal solutions for 14.5 degrees inclination mission when considering the spacecraft’s arrival mass at the Sun-Earth L4.

Spacecraft equipped with solar telescopes are planned for deployment at various vantage points in the heliosphere to conduct coordinated, multi-view observations of the Sun and its dynamic activities. We investigate solar visibility using imaging instruments aboard spacecraft stationed at the Sun-Earth Lagrange points L1, L4 and L5. First, the optimal arrival time for vertical periodic orbits at L4 and L5 is determined on the basis of geometric factors that maximize the visibility of the solar poles and higher latitudes. For different orbits around L1, L4 and L5, we calculate the visibility of the solar surface (i.e., observation days per year) as a function of solar latitude. Additionally, we analyze how the solar limb observed from one Lagrange point aligns with the solar surface visible from the other two points, with a focus on studying solar eruptions like flares and coronal mass ejections. This analysis aims to assess the feasibility of coordinated observations of off-disk erupting structures and their on-disk magnetic footpoints. Furthermore, we evaluate the improvement in continuous tracking of solar features, such as sunspots, with multiple spacecraft in various orbital configurations. This tracking helps in studying the long-term evolution of these features, from emergence to decay. A comprehensive comparison of observations from single (L1), double (L1 and L4), and multiple spacecraft (L1, L4 and L5) is conducted, aiding the design of future space missions involving solar observatories at the Sun-Earth Lagrange points.

The Sun-Earth Lagrange point L4 is considered as one of the unique places where the solar activity and heliospheric environment can be observed in a continuous and comprehensive manner. The L4 mission affords a clear and wide-angle view of the Sun-Earth line for the study of the Sun-Earth and Sun-Moon connections from he perspective of remote-sensing observations. In-situ measurements of the solar radiation, solar wind, and heliospheric magnetic field are critical components necessary for monitoring and forecasting the radiation environment as it relates to the issue of safe human exploration of the Moon and Mars. A dust detector on the ram side of the spacecraft allows for an unprecedented detection of local dust and its interactions with the heliosphere. The purpose of the present paper is to emphasize the importance of L4 observations as well as to outline a strategy for the planned L4 mission with remote and in-situ payloads onboard a Korean spacecraft. It is expected that the Korean L4 mission can significantly contribute to improving the space weather forecasting capability by enhancing the understanding of heliosphere through comprehensive and coordinated observations of the heliosphere at multi-points with other existing or planned L1 and L5 missions.

This paper introduces a procedure to optimize a low-thrust gravity-assist trajectory to the Earth–moon L1 periodic orbit utilizing the resonance-orbital structure as a guideline. The Earth–moon circular restricted three-body problem formulation is used to describe the problem. The proposed procedure determines the gravity-assist geometry and then finds the gravity-assist linking based on the multiple-point boundary value problem. The gravity-assist geometry determination step designs the periapsis rotation angle by solving a gradient descent optimization problem, yielding trajectories that break the symmetry of the resonance orbits. The multiple-point boundary-value problem seeks to solve a minimum-fuel problem linking two intermediate resonance-like orbits with rotated periapses. The first step of the optimal control problem establishes and solves a relatively easy two-point boundary problem approximating the original problem. The solution is used as the initial guess for the more complex multiple-point boundary value problem. The low-thrust resonance gravity-assist trajectory is compared to the trajectories designed based on traditional approaches involving low-thrust propulsion, demonstrating its validity and efficiency.

Space debris poses significant threats to spacecraft and human activities in space. Accurate modeling of space debris is crucial for understanding and mitigating these risks, ensuring the sustainability of the space environment. This paper discusses the importance of space debris modeling in the space environment, highlighting its critical role in safeguarding assets in orbit. Two primary methods of space debris modeling, namely the 1D and 3D approaches, are discussed in detail, and their respective strengths and limitations are elucidated. Furthermore, a comprehensive review of existing models, including the space debris evolutionary model (MOCAT, SOLEM, DAMAGE, LEODEEM & GEODEEM, DELTA, and LEGEND) and engineering models (MOCAT-MC, NEODEEM, MASTER, ORDEM), are presented. These models offer valuable insights into the dynamics and characteristics of space debris populations, aiding in formulating effective debris mitigation strategies and orbital capacity problems for reducing the possibilities of Kessler’s syndrome. Additionally, the paper provides insights into the ongoing development of the Korean space debris model, focusing on its methodology and space debris cataloging techniques for modeling space debris environments.

Recently, the number of space objects around the Earth has increased rapidly, necessitating systematic space risk management. This paper proposes a dual-mode framework for assessing the risk of collision between space objects. The proposed framework consists of microscopic and macroscopic modes. The former focuses on one-to-one collision events, and the latter assesses the overall collision risk inside a cell located in space. Two risk assessment case studies using the proposed two modes demonstrate the effectiveness of the proposed framework.

This paper presents the application of manifolds for space mission design in the Circular Restricted Three Body Problem, Bi-Circular Restricted Four Body Problem, and Patched-Planar Circular Restricted Three Body Problem. We tackle these three types of missions by defining each mission’s manifold-based transfer roadmap/sequence. The transfer to Sun–Earth L4 and L5 planar Lyapunov orbits utilizing the electric-propulsion-system-assisted artificial stable manifolds and the Sun–Earth system’s L1 and L2 stable and unstable manifolds are designed utilizing the circular restricted three-body problem. We also design the weak stability boundary transfers to Earth–Moon’s L1 and L2 utilizing Earth–Moon’s stable manifold and Sun–Earth’s L1 and L2 unstable manifolds utilizing the Bi-Circular Restricted Four Body Problem. We present a manifold projection method to quickly assess manifold reachability for multiple-moon tour analysis utilizing the patched-planar circular restricted three-body problem. A new filter is used during the Poincaré section analysis to determine optimal sets of stable and unstable manifolds for several types of missions. The multiple-point differential corrector is utilized with optimal manifolds to determine a continuous smooth trajectory for spacecraft with impulsive burn maneuvers.

This paper discusses the application of the homotopy approach to optimizing the free-flight-time low-thrust trajectory considering the performance characteristics. The trajectory optimization problem is formulated as an optimal control problem (two-point boundary value problem). The developed homotopy algorithm to solve the optimal control problem comprises three steps—orbital energy/acceleration matching, flight-time matching, and orbital reshaping. A comprehensive case study demonstrates the effectiveness of the proposed homotopy-based optimization algorithm for an optimal design of a low-thrust trajectory involving long mission time.