Astrodynamics最新文献

The classification of periodic relative motion in spacecraft dynamics is well-established, yet non-periodic motion remains underexplored. This paper addresses this gap by proposing a comprehensive classification model based on relative orbital dynamics. The elements of relative motion are defined, leveraging the characteristics of relative orbital dynamics to enable a detailed analysis of motion patterns. The analysis is then conducted from two perspectives: the geometrical configuration of the trajectory and the positional relationship between the trajectory and the spacecraft. A total of 19 motion styles are classified, including 14 non-periodic types, 3 periodic types, and 2 degenerate fixed-point cases, with corresponding dynamic conditions identified. Two representative applications demonstrate the practical utility of the classification: rapid target approach using flyby styles and hovering observation using droplet styles. Numerical simulations validate the accuracy of the classification and its applicability to real-world orbital scenarios. These findings provide a valuable framework for on-orbit services, such as formation flying, rendezvous, and orbital maintenance.

The growth in the overpopulation of resident space objects calls for space surveillance initiatives. In particular, the threat posed by in-orbit collisions and fragmentations, as well as by satellites re-entry requires an efficient space objects cataloguing capability. Ground-based sensors are the main contributors to build up and maintain a catalogue of space objects. In this context, survey radars can provide angular track, slant range, and Doppler shift measurements without the need for transit prediction, allowing either the refinement or the initial determination of the target orbital state. In the latter case, a proper Initial Orbit Determination (IOD) technique is required to reconstruct the orbital state of the observed object. This work presents the IODAD algorithm (Initial Orbit Determination from Angular and Doppler shift measurements), a novel radar IOD method when slant range is not available, and thus relying only on the angular and Doppler shift measurements. The proposed IOD algorithm combines the optical admissible region, computed from the angular track measurement, with the measured Doppler shift to compute a first estimate of the orbital state. This combination forks depending on whether the radar is monostatic or bistatic. At the end, the first estimate is refined through a batch filter and the IOD result is returned in terms of mean state and covariance. Unlike existing methods, the new algorithm offers greater flexibility and ease of operational application, as it does not need long measurements tracks as input, nor a specific advanced computational technique. Numerical simulations show the potential of the IODAD algorithm, both through nominal and sensitivity analysis, proving its validity to any survey radar. In addition, a comparison with an existing method demonstrates the significantly better performance of the proposed method. Finally, the results are confirmed by analysing a real dataset of transits concerning calibrator satellites.

This paper presents a methodology for determining the optimal transfer time in a kinetic impact defense strategy under uncertainties. Leveraging existing research, the paper first analyses the optimal deflection without considering uncertainty. It is then followed by an uncertainty-related analysis on the basis of the results. By mapping the angular parameters of the impact location, an uncertainty model is established and analytically propagated to the deflection model. The derivation of uncertainty propagation is verified with Monte Carlo simulation methods. Subsequently, the analytical expression for the deflection expectation under uncertainty is derived, and the optimal transfer time under uncertainty is investigated. The numerical results indicate that the research method proposed in this paper is effective, offering a robust approach for analyzing deflection missions under practical conditions.

The NASA Artemis program will take incremental steps towards establishing a sustainable and long-term human presence in the vicinity of the Moon. Artemis III and beyond (Artemis 3+) will leverage a mission architecture that supports science in lunar orbit and on the surface of the Moon. These later Artemis missions will exploit either the Gateway spacecraft, an orbital laboratory in a 9:2 lunar resonant Near-Rectilinear Halo Orbit (NRHO), or the Human Landing System (HLS) in a similar orbit. While the NRHO provides many benefits to operations, the multi-body gravitational regime produces complex dynamics that can make orbital transfers from Earth difficult to optimize. This study provides a comprehensive examination of NRHO rendezvous transfers subjected to operational constraints presented by an Artemis 3+ architecture. An ephemeris model is utilized for the outbound transfer to rendezvous with a 9:2 NRHO bounded by parameters, such as Earth parking orbit, ΔV requirements, and transfer time constraints expected from an Artemis 3+ mission. The model was used to compute ΔV optimal transfers to a rendezvous target at different epochs corresponding to different positions along the NRHO. Next, contour plots were generated for transfer ΔVs across a range of rendezvous epochs and cislunar transfer durations, similar to a porkchop plot. These contour plots allow mission designers to view the entire solution space over a selected window of dates instead of just point solutions for a given epoch. This is particularly useful for a problem that has multiple local minima. While there is already a theoretical understanding of optimal transfers to an NRHO, this work explores transfer and rendezvous optimization in the context of Artemis 3+ operational constraints. Contour plots of the transfer solution space generated in this study are shown to be a valuable tool for constructing continuous and optimal launch windows.

While the exponential growth of the space sector and new operative concepts ask for higher spacecraft autonomy, the development of AI-assisted space systems was so far hindered by the low availability of power and energy typical of space applications. In this context, Spiking Neural Networks (SNN) are highly attractive because of their theoretically superior energy efficiency due to their inherently sparse activity induced by neurons communicating by means of binary spikes. Nevertheless, the ability of SNN to reach such efficiency on real world tasks is still to be demonstrated in practice. To evaluate the feasibility of utilizing SNN onboard spacecraft, this work presents a numerical analysis and comparison of different SNN techniques applied to scene classification for the EuroSAT dataset. Such tasks are of primary importance for space applications and constitute a valuable test case given the abundance of competitive methods available to establish a benchmark. Particular emphasis is placed on models based on temporal coding, where crucial information is encoded in the timing of neuron spikes. These models promise even greater efficiency of resulting networks, as they maximize the sparsity properties inherent in SNN. A reliable metric capable of comparing different architectures in a hardware-agnostic way is developed to establish a clear theoretical dependence between architecture parameters and the energy consumption that can be expected onboard the spacecraft. The potential of this novel method and its flexibility to describe specific hardware platforms is demonstrated by its application to predicting the energy consumption of a BrainChip Akida AKD1000 neuromorphic processor.

A long time series prediction of hypersonic glide vehicles (HGVs) is urgently needed to fulfill the increasing defense demand since trajectory prediction is crucial for fight intention inference, threat assessment, and vehicle interception in near space. To address this issue, a trajectory prediction algorithm based on maneuver analysis under a neural network is proposed for HGVs. From a dynamic perspective, trajectory prediction can be mapped as a time series prediction by maneuvering parameters. First, a new set of maneuvering parameters is extracted from the maneuvering mode analysis, which aids the fast Fourier transform (FFT) method in the trajectory classification implementation. Subsequently, the equilibrium glide trajectory can be forecasted by employing a novel trajectory modeling algorithm based on the autoregression method. Moreover, the trajectory with a large number of maneuvers that exceeds the proper threshold value is addressed via the error compensation method and the intelligent prediction algorithm. The error compensation method eliminates the disadvantages arising from the limited window length but low frequency of the trajectory data under periodic variation and the anisotropy of the prediction capacity in different dimensions. Finally, the advanced self-attention and distilling mechanism and generative style encoder implementations are introduced to improve the prediction capacity in the Transformer framework, which ensures the time sensitivity of the improved prediction algorithm to meet the defense purpose. The extensive flight scenarios of the HGVs demonstrate that the novel algorithm outperforms the existing trajectory prediction algorithms.

Liquid hydrogen (LH₂) is a highly preferred fuel for propelling large-ton Martian landers. Currently, composite insulation structures (CISs), consisting of multilayer insulation (MLI) and spray-on foam insulation (SOFI) are used for the thermal protection of cryogenic propellant tanks. However, the Martian environment poses significant challenges for the thermal protection of LH₂ tanks, primarily because of the presence of rarefied carbon dioxide (CO₂). CO₂ can condense within the MLI microchannels, leading to significant deterioration in the insulation performance of CIS. This study investigates this unique phenomenon occurring on Mars. Two thermal-insulation schemes are designed for potential applications in Martian landers. Numerical models are developed to predict the insulation performance of CIS on Mars by coupling CO₂ condensation and conduction calculations within microchannels. Several key findings emerge: (1) The heat leakage decreases by 82.22% as SOFI thickness increases from 5 to 50 mm. (2) The heat leakage initially rises sharply, and then increases gradually with ambient Martian temperature. (3) The insulation performance deteriorates rapidly and then stabilizes as the number of MLI layers increases, but the heat leakage of CIS exceeds 10 W/m². (4) A new type of insulation scheme (NIS) aimed at solving the issue in finding (3) is designed for the Mars lander by prioritizing the minimum weight which results in an NIS heat leakage of 0.298 W/m²—two to four orders of magnitude lower than that of CISs. These findings provide valuable technical insights into enhancing the thermal protection of LH₂ propellant tanks.

This paper proposes a distributed composite state observer-based cooperative control scheme for multiple spacecraft with consideration of external disturbances. The nonlinear relative motion dynamics of multiple spacecraft is established in a rotating local-vertical local-horizontal (LVLH) reference frame attached to a virtual spacecraft. Based on the undirected topology, a novel distributed composite state observer is developed for each follower spacecraft only using the position level information of its neighbors, such that the states and total disturbances of both the leader spacecraft dynamical model and the consistency error system can be estimated. Then, a composite controller is put forward in terms of the outputs of the developed composite state observer, which not only enables to overcome the external disturbances but also achieve spacecraft consistent collaborative control or formation flying. The convergence and stability of the distributed composite state observer and the closed-loop control system are verified by the Lyapunov theory. Finally, the effectiveness of the presented cooperative control strategy is proved by two illustrative numerical simulations.

The Earth-Moon L₁, L₂ halo orbit families support a variety of options for lunar surface activities as well as other developments in the cislunar region. The planned operational orbit for NASA’s Gateway is the 9:2 L₂ synodic resonant halo orbit. Furthermore, orbits from this family could serve as staging locations for future missions to near-Earth asteroids and Mars, hubs for servicing logistics, and nodes for sensor networks. Therefore, it is essential to understand the dynamics that govern the halo orbit regions. Previously, a sub-region in the L₂ halo family of the Earth-Moon system, termed the “Interface Region” was identified to be particularly challenging when transitioning the solutions from this L₂ halo family sub-region to the Higher-Fidelity Ephemeris model (HFEM). Recent research suggests that the intermediate models incorporating pulsation in the Earth-Moon motion offer more insight into the dynamics governing this general underlying flow. This investigation aims to characterize the interface region by understanding the dynamical structures via the Hill Restricted Four-Body Problem (HR4BP).