Astrodynamics最新文献

The regularization theory has successfully enabled the removal of gravitational singularities associated with celestial bodies. In this study, regularizing techniques are merged into a multi-impulse trajectory design framework that requires delicate computations, particularly for a fuel minimization problem. Regularized variables based on the Levi–Civita or Kustaanheimo–Stiefel transformations express instantaneous velocity changes in a gradient-based direct optimization method. The formulation removes the adverse singularities associated with the null thrust impulses from the derivatives of an objective function in the fuel minimization problem. The favorite singularity-free property enables the accurate reduction of unnecessary impulses and the generation of necessary impulses for local optimal solutions in an automatic manner. Examples of fuel-optimal multi-impulse trajectories are presented, including novel transfer solutions between a near-rectilinear halo orbit and a distant retrograde orbit.

Spinning electrodynamic tether systems (SEDTs) have promising potential for the active removal of space debris, the construction of observation platforms, and the formation of artificial gravity. However, owing to the survivability problem of long tethers, designing collision-avoidance strategies for SEDTs with space debris is an urgent issue. This study focuses on the design of collision-avoidance strategies for SEDTs with an electrodynamic force (Ampere force). The relative distance between the debris and the SEDT is first derived, and then two collision-avoidance strategies are proposed according to the two different cases. When debris collides closer to a lighter subsatellite, a stationary avoidance strategy is proposed to change the spatial position of the subsatellite by adjusting only the angular motion of the tether, which maintains the original orbit of the SEDT. When debris collides closer to a heavier main spacecraft, a comprehensive avoidance strategy is proposed to change the spatial position of the SEDT by slightly modifying the orbital height and changing the tether angular motion simultaneously. The numerical results illustrate that the proposed strategies promptly avoid potential collisions of an SEDT with space debris without significant changes in the orbital parameters of the SEDT.

This study focuses on stabilizing the libration dynamics of an electrodynamic tether system (EDTS) using generalized torques induced by the Lorentz force. In contrast to existing numerical optimization methods, a novel analytical feedback control law is developed to stabilize the in-plane and out-of-plane motions of a tether by modulating the electric current only. The saturation constraint on the current is accounted for by adding an auxiliary dynamic system to the EDTS. To enhance the robustness of the proposed controller, multiple perturbations of the orbital dynamics, modeling uncertainties, and external disturbances are approximated using a neural network in which the weighting matrix and approximation error are estimated simultaneously, such that these perturbations are well compensated for during the control design of the EDTS. Furthermore, a dynamically scaled generalized inverse is utilized to address the singular matrix in the control law. The closed-loop system is proven to be ultimately bounded based on Lyapunov stability theory. Finally, numerical simulations are performed to demonstrate the effectiveness of the proposed analytical control law.

This study proposes a novel adaptive neural dynamic-based hybrid control strategy for stable subsatellite retrieval of two-body tethered satellite systems. The retrieval speed is given analytically, ensuring a libration-free steady state. To mitigate the potential libration motion, a general control input signal is generated by an adaptive neural-dynamic (AND) algorithm and executed by adjusting the retrieval speed and thruster on the subsatellite. To address the limited retrieval speed and improve the control performance, the thruster controller is manipulated according to a novel advanced state fuzzy control law based on higher-order libration states, whereas the remaining control input is allocated to the speed controller. The Lyapunov stability of the control strategy is demonstrated analytically. Numerical simulations validate the proposed control strategy, demonstrating well-allocated control inputs for both controllers and good control performance.

An algorithm for robust initial orbit determination (IOD) under perturbed orbital dynamics is presented. By leveraging map inversion techniques defined in the algebra of Taylor polynomials, this tool returns a highly accurate solution to the IOD problem and estimates a range centered on the aforementioned solution in which the true orbit should lie. To meet the specified accuracy requirements, automatic domain splitting is used to wrap the IOD routines and ensure that the local truncation error, introduced by a polynomial representation of the state estimate, remains below a predefined threshold. The algorithm is presented for three types of ground-based sensors, namely range radars, Doppler-only radars, and optical telescopes, by considering their different constraints in terms of available measurements and sensor noise. Finally, the improvement in performance with respect to a Keplerian-based IOD solution is demonstrated using large-scale numerical simulations over a subset of tracked objects in low Earth orbit.

This study involved simulations and experiments aimed at assessing the efficacy of a tether net in encapsulating space debris. The tether net was modeled as a spring–mass–damper system considering the influence of aerodynamic and gravitational forces and the occurrence of debris collisions. To examine the influence of collision position and size disparity between the debris and the net on debris capture status, the entanglement nodes of the net were identified. Experiments were conducted to evaluate the wrapping capabilities of the tether net, focusing specifically on debris capture. Subsequently, the results were compared with those of the numerical simulation. In the experiments, radio frequency identification was used to identify the entanglement points of the tether net. Previous studies have indicated that the ideal collision point for capturing debris using a tether net with the debris intended to be captured is located at the center of the net. However, the experimental results of this study revealed that a collision position that is slightly shifted from the center of the tether net is more advantageous for capturing debris in terms of tether net entanglement.

This study aims to assess the autonomous navigation performance of an asteroid orbiter enhanced using an inter-satellite link to a beacon satellite. Autonomous navigation includes the orbit determination of the orbiter and beacon, and asteroid gravity estimation without any ground station support. Navigation measurements were acquired using satellite-to-satellite tracking (SST) and optical observation of asteroid surface landmarks. This study presents a new orbiter–beacon SST scheme, in which the orbiter circumnavigates the asteroid in a low-altitude strongly-perturbed orbit, and the beacon remains in a high-altitude weakly-perturbed orbit. We used Asteroid 433 Eros as an example, and analyzed and designed low- and high-altitude orbits for the orbiter and beacon. The navigation measurements were precisely modeled, extended Kalman filters were devised, and observation configuration was analyzed using the Cramer–Rao lower bound (CRLB). Monte Carlo simulations were carried out to assess the effects of the orbital inclination and altitudes of the orbiter and beacon as key influencing factors. The simulation results showed that the proposed SST scheme was an effective solution for enhancing the autonomous navigation performance of the orbiter, particularly for improving the accuracy of gravity estimation.

Large spacecraft fall out of orbit and re-enter the atmosphere at the end of their lifetime, and they can break up into small debris upon re-entry. The spacecraft debris generated by the disintegration may lead to high risk when the surviving debris reaches the ground. One way to reduce the damage risk of spacecraft is to simulate the spacecraft disintegration process and accurately predict the falling area. Aerodynamics seriously affects the reentering process, especially in the continuous flow regime. Aerodynamic force and heat are the main factors leading to debris disintegration. High dynamic pressure leads to sharp changes in attitude and complex trajectories during debris fall. A numerical method based on an unstructured Cartesian grid was developed to simulate the disintegrating separation problem by coupling the Navier-Stokes equation and the six-degree-of-freedom trajectory equation. A method combining the numerical method for dynamic processes with numerical simulation based on a static aerodynamic/dynamic characteristic database was developed for forecasting the falling area. Spacecraft disintegrating separation from 60 km was simulated using the method, and the multibody aerodynamic interference and the separation trajectory were predicted. The falling process was forecast by a numerical simulation method based on the static aerodynamic database/dynamic characteristic database when the debris went out of the influence domain. This method has good forecasting efficiency while considering the aerodynamic interference, making it a valuable method for forecasting disintegrating separation and falling debris.

Currently, a surge in the number of spacecraft and fragments is observed, leading to more frequent breakup events in low Earth orbits (LEOs). The causes of these events are being identified, and specific triggers, such as collisions or explosions, are being examined for their importance to space traffic management. Backward propagation methods were employed to trace the origins of these types of breakup events. Simulations were conducted using the NASA standard breakup model, and satellite Hitomi’s breakup was analyzed. Kullback-Leibler (KL) divergences, Euclidean 2-norms, and Jensen-Shannon (JS) divergences were computed to deduce potential types of breakups and the associated fragmentation masses. In the simulated case, a discrepancy of 22.12 s between the estimated and actual time was noted. Additionally, the breakup of the Hitomi satellite was estimated to have occurred around UTC 1:49:26.4 on March 26, 2016. This contrasts with the epoch provided by the Joint Space Operation Center, which was estimated to be at 1:42 UTC ± 11 min. From the findings, it was suggested that the techniques introduced in the study can be effectively used to trace the origins of short-term breakup events and to deduce the types of collisions and fragmentation masses under certain conditions.