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Orbit accuracy of the transfer orbit and the mission orbit is the basis for the orbit control of all-electric-propulsion Geostationary Orbit (GEO) satellites. Global Navigation Satellite System (GNSS) simulation data are used to analyze the main factors affecting GEO satellite orbit prediction accuracy under the no-thrust condition, and an electric propulsion calibration algorithm is designed to analyze the orbit determination and prediction accuracy under the thrust condition. The calculation results show that the orbit determination accuracy of mission orbit and transfer orbit without thrust is better than 10 m using onboard GNSS technology. The calibration accuracy of electric thrust is about 10−9 m/s2 and 10−7 m/s2 with 40 h and 16 h arc length, respectively, using the satellite self-positioning data of 100 m accuracy to calibrate the electric thrust. If satellite self-positioning data accuracy is at the 10 m level, the electric thrust calibration accuracy can be improved by about one order of magnitude, and the 14-day prediction accuracy of the transfer orbit with thrust is better than 1 km.

Midcourse correction design is key to space transfers in the cislunar space. Autonomous guidance has garnered significant attention for its promise to decrease the dependence on ground control systems. This study addresses the problem of midcourse corrections for Earth-Moon transfer orbits based on high-order state transition tensors (STTs). The scenarios considered are direct Earth-Moon transfers and low-energy transfers to lunar distant retrograde orbits (DROs), where the latter involve weak stability boundary (WSB) and lunar gravity assist (LGA) techniques. Semi-analytical formulas are provided for computing the trajectory correction maneuvers (TCMs) using high-order STTs derived using the differential algebraic method. Monte Carlo simulations are performed to evaluate the effectiveness of the proposed approach. Compared with existing explicit guidance algorithms, the STT-based approach is much cheaper computationally and features fewer final position errors. These results are promising for fast and efficient orbital autonomous correction guidance approaches in the cislunar space.

In this paper, we study the long-term dynamical evolution of highly elliptical orbits in the medium-Earth orbit region around the Earth. The real population consists primarily of Geosynchronous Transfer Orbits (GTOs), launched at specific inclinations, Molniya-type satellites and related debris. We performed a suite of long-term numerical integrations (up to 200 years) within a realistic dynamical model, aimed primarily at recording the dynamical lifetime of such orbits (defined as the time needed for atmospheric reentry) and understanding its dependence on initial conditions and other parameters, such as the area-to-mass ratio (A / m). Our results are presented in the form of 2-D lifetime maps, for different values of inclination, A / m, and drag coefficient. We find that the majority of small debris (\\[>70\\%\\], depending on the inclination) can naturally reenter within 25–90 years, but these numbers are significantly less optimistic for large debris (e.g., upper stages), with the notable exception of those launched from high latitude (Baikonur). We estimate the reentry probability and mean dynamical lifetime for different classes of GTOs and we find that both quantities depend primarily and strongly on initial perigee altitude. Atmospheric drag and higher A / m values extend the reentry zones, especially at low inclinations. For high inclinations, this dependence is weakened, as the primary mechanisms leading to reentry are overlapping lunisolar resonances. This study forms part of the EC-funded (H2020) “ReDSHIFT” project.

In response to the increasing demand for high-precision navigation of satellites operating in the cislunar space, this study introduces an onboard orbit determination algorithm considering both convergence and computational efficiency, referred to as the Sliding-Window Batch Processing (SWBP) algorithm. This algorithm combines the strengths of data batch processing and the sequential processing algorithm, utilizing measurement data from multiple historical and current epochs to update the orbit state of the current epoch. This algorithm facilitates rapid convergence in orbit determination, even in instances where the initial orbit error is large. The SWBP algorithm has been used to evaluate the navigation performance in the Distant Retrograde Orbit (DRO) and the Earth–Moon transfer orbit. The scenario involves a low-Earth-orbit (LEO) satellite establishing satellite-to-satellite tracking (SST) links with both a DRO satellite and an Earth–Moon transfer satellite. The LEO satellite can determine its orbit accurately by receiving GNSS signals. The experiments show that the DRO satellite achieves an orbit determination accuracy of 100 m within 100 h under an initial position error of 500 km, and the transfer orbit satellite reaches an orbit determination accuracy of 600 m within 3.5 h under an initial position error of 100 km. When the Earth–Moon transfer satellite exhibits a large initial orbital error (on the order of hundreds of kilometers) or the LEO satellite’s positional accuracy is degraded, the SWBP algorithm demonstrates superior convergence speed and precision in orbit determination compared to the Extended Kalman Filter (EKF). This confirms the proposed algorithm’s capability to handle complex orbital determination scenarios effectively.

Rideshares increase launch capabilities and decrease the launch costs. However, the range of orbits available for secondary payloads is dependent on launch constraints for the primary mission. Additionally, communications constraints and limited propellant options must be incorporated in the preliminary mission design strategy for secondary payloads. Ridesharing opportunities are now available for orbit destinations beyond Low Earth Orbit (LEO). In this investigation, transfers from Geosynchronous Transfer Orbits (GTO) to Sun-Earth libration point orbits are constructed by leveraging stable manifold structures and Poincaré maps.

There are plans to set up a space-based gravitational wave observatory that will use an ultra-large-scale laser interferometer in space to detect medium- and low-frequency gravitational waves. Both heliocentric and geocentric formations adopt the method of launching three satellites with one rocket, which has high requirements in terms of the carrying capacity of the rocket. Therefore, a proper transfer design is a prerequisite for achieving space-based gravitational wave detection. In this paper, the transfer orbit for three satellites of the Taiji mission is designed based on the two-impulse transfer model. Moreover, the influence on orbit design of the position of the formation relative to Earth, the initial phase angle of the formation, and the initial time of transfer is analyzed. The Earth-leading and -trailing transfers show opposite patterns in the above three aspects. A smaller velocity increment is required if a proper initial time is selected. After taking into account the stability of the formation, C3, the required velocity increment, transfer time, and the distance to Earth, 20° is determined to be the optimal initial trailing/leading angle.

We present a study of the transfer orbits connecting landing points of irregular-shaped asteroids. The landing points do not touch the surface of the asteroids and are chosen several meters above the surface. The ant colony optimization technique is used to calculate the multiple-hopping trajectories near an arbitrary irregular asteroid. This new method has three steps which are as follows: (1) the search of the maximal clique of candidate target landing points; (2) leg optimization connecting all landing point pairs; and (3) the hopping sequence optimization. In particular this method is applied to asteroids 433 Eros and 216 Kleopatra. We impose a critical constraint on the target landing points to allow for extensive exploration of the asteroid: the relative distance between all the arrived target positions should be larger than a minimum allowed value. Ant colony optimization is applied to find the set and sequence of targets, and the differential evolution algorithm is used to solve for the hopping orbits. The minimum-velocity increment tours of hopping trajectories connecting all the landing positions are obtained by ant colony optimization. The results from different size asteroids indicate that the cost of the minimum velocity-increment tour depends on the size of the asteroids.

In this paper, the composition of the drag-free CubeSat platform was introduced first. Then the control strategy of drag-free CubeSat orbit keeping was proposed. Finally, the simulation with the control scheme of the Homan transfer orbit was compared. By comparing and analyzing the fuel consumption of the drag-free track and Homan transfer track, it is demonstrated that the continuous drag compensation of the drag-free control scheme can significantly reduce the consumption of thruster mass, and the drag-free technology has good advantages in track stability.

The problem of finding a planar two-impulse transfer orbit between two known elliptical orbits that minimizes the total characteristic velocity of the transfer arc is examined. Using a transformation of variables presented in previous work, necessary conditions for an optimal transfer are determined, followed by a proof that an optimal transfer exists. We then consider the problem of finding a minimizing planar two-impulse transfer over the set of two-impulse transverse transfers. A minimizing solution for this problem requires that either each of the boundary orbits has an apse that is the same distance from the center of attraction as the other, or else the boundary orbits are coaxial. The transfer orbits are tangent to the boundary orbits at apses. Minimizing solutions of the transverse transfer problem are found in closed form.

Near Earth Asteroids pose significant risks of potential impacts on Earth, while simultaneously offering valuable opportunities for asteroid capturing missions. This study investigated the possibility of using distant retrograde orbits (DROs) in the Earth-Moon system as a parking orbit for asteroids capturing missions. Previous research has mainly focused on capturing asteroids into unstable orbits like Lyapunov orbits where no maneuver is needed inserting into these orbits. However, the stability of DROs makes the previous methods not suitable and rises the difficulty to design a low-energy transfer to DROs. To address the limitation, this paper proposed an approach for identifying transfer behaviours around DROs by developing the applications of Finite-Time Lyapunov Exponent (FTLE) fields. By applying the method of FTLE fields to identify the transfer behaviours around DROs, efficient insertion angles for DROs are achieved and low-energy transfer trajectories to DROs are acquired. Numerical simulations indicate that the insertion speed increment of DROs is less than the station-keeping speed increment of Lyapunov orbits. Additionally, to expand the application of captures asteroids, this study investigated the feasibility and efficiency of using captured asteroids as an enhanced kinetic impactor for asteroid defence missions.