Spacecraft Maneuvers Research Articles

Optimization of tethered artificial gravity assists for capture about binary asteroids in the circular restricted three-body problem

The use of tethered artificial gravity assists in space travel to perform orbital maneuvers has the potential to significantly extend missions through the reduction of fuel consumption. As space exploration in the solar system has become more accessible, interest in near-Earth asteroid (NEA) missions has grown. Studying NEAs can yield new information about the early solar system as well as provide important insights into related fields, such as planetary defense. Further, it has been discovered that many NEAs are in binary systems, which presents a unique orbital environment for optimizing spacecraft maneuvers for future asteroid missions. This paper develops an approach employing a genetic algorithm to determine optimal tethered maneuvers to periodic orbits in the vicinity of a binary asteroid using circular restricted three-body problem (CR3BP) dynamics. The optimization identifies the maneuver that results in the greatest change in the Jacobi constant values between the incoming orbit and final desired orbit. This maximizes the benefit of the tethered artificial gravity assist into a periodic orbit about the primary asteroid. The tethered gravity assists are examined for binary asteroid systems with various mass ratios using CR3BP dynamics. This investigation selected three observed binary asteroid systems with orbital eccentricities that allow for modelling using the CR3BP. Optimizations were performed and simulated for the three systems with different mass ratios to investigate and demonstrate the viability of tethered gravity assists in various binary asteroid environments.

In-space cybernetical intelligence perspective on informatics, manufacturing and integrated control for the space exploration industry

The integration of cybernetic principles into space technology has led to a significant shift in spacecraft guidance systems and space station operations. This scholarly work provides a comprehensive analysis of the profound impact that cybernetics has had on space technology, particularly focusing on the development and implementation of closed-loop control systems. Building on foundational contributions to cybernetics, this paper offers a detailed analysis of the complex interplay between control mechanisms, behavioral dynamics, and information management in space informatics. The essential role of cybernetic control systems is highlighted through their critical function in enabling precise spacecraft maneuvers, as demonstrated by the guidance systems used in Apollo lunar missions during the crucial descent and landing phases. In modern space station programs, cybernetic intelligence operates much like a neural network, processing data in real-time and coordinating the various elements of the station's infrastructure. This cybernetic system excels in a wide range of tasks, from space docking maneuvers to waste management, thereby safeguarding astronaut welfare and ensuring mission success. Information serves as the cornerstone of problem analysis and processing, providing a holistic understanding of information flow within complex systems. The study concludes by highlighting the application of a cybernetically intelligent neural network in developing adaptive learning systems for nonlinear and intricate industrial processes. Ultimately, this paper underscores the transformative role of cybernetics in shaping the future of space exploration, emphasizing the seamless integration of control, information, and behavior in space technology to advance our understanding of celestial systems while enhancing the efficiency and safety of space missions.

Nanosatellite Design Considerations for a Mission to Explore the Propellant Sloshing Problem

Sloshing Platform for In-Orbit Controller Experimentation is an ambitious, student run mission to design and fly a cubesat to study fluid sloshing in spacecraft. The project will examine zero-g propellant sloshing from an experimental standpoint. Despite the small size and limited payload capacity, we intend to use the cubesat platform to mimic larger spacecraft and implement novel detection and computer vision methods in our analysis. Many modern spacecraft rely on propellant-filled tanks to perform attitude control and station-keeping maneuvers. When a large percentage of the spacecraft’s mass is comprised of liquid propellant, sloshing becomes a critical aspect of spacecraft attitude control and stability. The mission will study the tank/fluid dynamics using new methods to gain an enhanced understanding of low-gravity fluid disturbance effects and improve simulations using equivalent mechanical models (EMMs). Active control of the fluid leading to the reduction of propellant slosh settling times will improve the maneuverability and performance of spacecraft. This paper will focus on satellite payload research and design requirements used to inform other aspects of the SPICEsat design. In this paper, mission objectives will be discussed, numerical simulations for the proposed control algorithms are demonstrated, and a satellite experiment design is presented. Finally, we examine computational fluid dynamics models to validate the satellite design and propellant sensing components of the proposed spacecraft.

Observation Method for Autonomous Maneuver of Spacecraft under Emergency Conditions

Observation Method for Autonomous Maneuver of Spacecraft under Emergency Conditions

Research on relative reachable domain in target orbit for maneuvering spacecraft

PurposeLots of successful space missions require that the maneuvering spacecraft can reach the target spacecraft. Therefore, research on relative reachable domain (RRD) in target orbit for maneuvering spacecraft is particularly important and is currently a hot-debated topic in the field of aerospace. This paper aims at analyzing and simulating the RRD in target orbit for maneuvering spacecrafts with a single fixed-magnitude impulse and continuous thrust, respectively, to provide a basis for analyzing the feasibility of spacecraft maneuvering missions and improving the design efficiency of spacecraft maneuvering missions.Design/methodology/approachBased on the kinematics model of relative motion, RRD in target orbit for maneuvering spacecraft with a single fixed-magnitude impulse can be calculated via analyzing the relationship between orbital elements, position vector and velocity vector of spacecrafts, and relevant studies are introduced to compare simulation results for the same case and validate the method proposed in the paper. With analysis of the dynamic model of relative motion, the calculation of RRD in target orbit for maneuvering spacecraft with continuous thrust can be transformed as the solution of the optimal control problem, and example emulations are carried out to validate the method.FindingsFor the case with a single fixed-magnitude impulse, simulation results show preliminarily that the method is in agreement with the method in Ref. (Wen et al., 2016), which treats the same case and thus is plausibly correct and feasible. For the case with continuous thrust, analysis and simulation results confirm the validity of the proposed method. The methods based on relative motion in this paper can efficiently determining the RRD in target orbit for maneuvering spacecraft.Originality/valueBoth theoretical analyses and simulation results indicate that the method proposed in this paper is comparatively simple but efficient for determine the RRD in target orbit for maneuvering spacecraft swiftly and precisely.

Responsive Maneuver Planning for Sun-Synchronous Repeating Ground Track Orbits

This paper investigates the suitability of sun-synchronous orbits (SSOs) as an injection location for responsive, maneuverable spacecraft. This work identifies a range of near-circular, repeating ground track (RGT) orbits that can be easily maneuvered into from SSOs. Four orbits that are simultaneously sun-synchronous and have an RGT are found, which could deliver high-frequency imagery of a target with constant lighting conditions. A method for maneuvering between RGT orbits that are targeting different regions using high-thrust impulse propulsion is also presented. This method is used to model two possible case studies for targeting regions on demand starting from an SSO injection orbit. In a comparative case study with a method that used low-thrust propulsion to move between targets, it was found that the method outlined in this paper required 55% less ΔV. A case study considering urgent response times showed that RGTs covering targets could be reached in as little as 4 h from any starting location, but this could require extremely large amounts of ΔV (upward of 1 km/s), while any global target could be viewed within 18 h with a maximum ΔV of 122.5 m/s.

The ReSWARM microgravity flight experiments: Planning, control, and model estimation for on‐orbit close proximity operations

AbstractOn‐orbit close proximity operations involve robotic spacecraft maneuvering and making decisions for a growing number of mission scenarios demanding autonomy, including on‐orbit assembly, repair, and astronaut assistance. Of these scenarios, on‐orbit assembly is an enabling technology that will allow large space structures to be built in situ, using smaller building block modules. However, like many of these scenarios, robotic on‐orbit assembly involves several technical hurdles, such as changing system models. For instance, grappled modules moved by a free‐flying “assembler” robot can cause significant changes in the combined system inertia, which have cascading impacts on motion planning and control portions of the autonomy stack. Further, on‐orbit assembly and other scenarios require collision‐avoiding motion planning, particularly when operating in a “construction site” scenario of multiple assembler robots and structures. Multiple key technologies that address these complicating factors for autonomous microgravity close proximity operations are detailed in this work, in particular: (1) application of global long‐horizon planning, accomplished using offline and online sampling‐based planner options that consider the system dynamics; (2) adaptation of the recently proposed RATTLE information‐aware planning framework for on‐orbit reconfiguration model learning; and (3) connection with robust control tools to provide low‐level control robustness using current system knowledge. These approaches were demonstrated for an autonomous on‐orbit assembly use case by the RElative Satellite sWarming and Robotic Maneuvering (ReSWARM) experiments using NASA's Astrobee robots on the International Space Station. Results of the ReSWARM experiments are provided along with significant operational and implementation detail discussing the practicalities of hardware implementation and unique aspects of working with the Astrobee free‐flyer robots in microgravity. ReSWARM provides a base set of planning and control tools for robotic close proximity operations, demonstrates them in microgravity, and outlines some of the important hardware aspects that future autonomous free‐flyers will need to consider.

A Bayesian Approach to Low-Thrust Maneuvering Spacecraft Tracking

Bayesian estimation with an explicit transitional prior is required for a tracking algorithm to be embedded in most multitarget tracking frameworks. This paper describes a novel approach capable of tracking maneuvering spacecraft with an explicit transitional prior and in a Bayesian framework, with fewer than two observations passes per day. The algorithm samples thrust profiles according to a multivariate Laplace distribution. It is shown that multivariate Laplace distributions are particularly suited to track maneuvering spacecraft, leading to a log probability function that is almost linear with the thrust. Principles from rare event simulation theory are used to propagate the tails of the distribution. Fast propagation is enabled by multi-fidelity methods. Because of the diffuse transitional prior, a novel k-nearest-neighbor-based ensemble Gaussian mixture filter is developed and used. The method allows Bayesian tracking of maneuvering spacecraft for several scenarios with fewer than two measurement passes per day and with a mismatch between the true and expected thrust magnitude of up to a factor of 200. The validity domain and statistical significance of the method are shown by simulation through several Monte Carlo trials in different scenarios and with different filter settings.

Anti-unwinding adaptive robust backstepping attitude maneuver control for rigid spacecraft

In this paper, an anti-unwinding adaptive robust backstepping control law is proposed to solve the attitude control problem of rigid spacecraft with external disturbances and inertia uncertainties. A new variable based on hyperbolic sine functions is designed to prevent the unwinding phenomenon, and then a novel robust backstepping controller is developed. Initially, a nonlinear tracking law is designed to obtain the desired angular velocity and to ensure that the attitude maneuvering system could have two equilibrium points. Subsequently, an adaptive robust control law is devised, which incorporates the adaptive estimation method to update the inertial parameters in real-time. Moreover, the bounded external disturbances are taken into account in the design of the control law, and thus the robustness of the attitude control system is improved. Furthermore, the global asymptotic stability of the attitude control system is verified by the Lyapunov-Krasovskii theorem. Finally, numerical simulations are carried out to demonstrate the effectiveness and robustness of the proposed control law, as well as the anti-unwinding characteristics are perfectly validated during the attitude maneuver of rigid spacecraft.

Detection and estimation of spacecraft maneuvers for catalog maintenance

Building and maintaining a catalog of resident space objects involves several tasks, ranging from observations to data analysis. Once acquired, the knowledge of a space object needs to be updated following a dedicated observing schedule. Dynamics mismodeling and unknown maneuvers can alter the catalog’s accuracy, resulting in uncorrelated observations originating from the same object. Starting from two independent orbits, this work presents a convex-optimization based approach to find admissible maneuvers to correlate the two orbits. In case of a feasible maneuver detection, the time profile of the maneuver is estimated, exploiting the conjugate unscented transform to map the orbits’ covariance into the maneuver uncertainty. It is also shown that a careful choice of the coordinate system is pivotal to obtain the solution in one iteration, without successive convexification.

Cascade nonlinear observer design for large amplitude of slosh variables in spacecraft maneuver with liquid propellant tank

Slosh dynamics in a liquid propellant tank during a spacecraft maneuver can have an effect on the system stability and potentially lead to critical problems for the mission operation. In particular, the fuel sloshing problem has drawn significant attention in aerospace applications where a large portion of the gross weight is comprised of liquid fuel. Recently, many active control methods have been introduced as alternatives to passive approaches. However, the majority of these methods basically require the full-state condition. To guarantee the availability of the full-state condition, this paper proposes a cascade nonlinear state observer based on the sliding mode theory for estimating the unmeasurable slosh variables in a planar model. Additionally, for large amplitude estimation, a complete nonlinear mathematical model with a pendulum analogy is formulated without a linearization process. The stability analysis is theoretically proven by Lyapunov method. The effectiveness is demonstrated through simulations, and the performance comparison with a different observer method is also presented in the results.

GEO spacecraft maneuver detection based on causal inference

Geosynchronous (GEO) spacecraft maneuver detection is deemed crucial for space domain awareness (SDA) as it helps to maintain catalogs of high value resident space objects (RSOs). However, several difficulties in resolving maneuver time and requirements for historical data quality still exist for the current research. This paper proposes a method based on causal inference for accurate and timely detection of the maneuver using real-time observational data from optical sensors. Firstly, the residual information of time series observations is obtained through counterfactual reasoning. Secondly, according to the Gaussianity of the optical sensor observation noise, a structural causal equation (SCE) of the residuals of time series observations is derived. Lastly, the maneuver is abstracted as a soft intervention, and the minimum average treatment effect (MATE) is proposed to estimate the effect of the maneuver and used as a feature parameter to identify the GEO spacecraft maneuver. After experimental analysis, the detection rate and detection timeliness of the proposed method reach more than 90%, a significant improvement compared with the conventional χ2, skewness, and kurtosis test methods.

Обзор маневров космических аппаратов на околокруговых орбитах

This paper presents comprehensive information about spacecraft (SC) maneuvers in the vicinity of a circular orbit. The implementation of SC maneuvers in the vicinity of circular orbits makes it possible to solve many research and applied problems. Typical examples are the problems of rendezvous and docking of SC moving in near-circular orbits, the implementation of a group flight of several SC in a given configuration, the deployment and maintenance of low-orbit systems to solve the problems of providing communications and remote sensing of the Earth. Practical examples of SC maneuvers are given.

Nonlinear attitude maneuvering of a flexible spacecraft for space debris tracking and collision avoidance

This work addresses the problem of attitude maneuvering of a spacecraft equipped with two flexible solar panels, in response to an alert in two distinct operational scenarios: (a) detection of an approaching debris, and (b) collision avoidance of an impacting debris. In scenario (a) the approaching debris is assumed to be off a collision course with the satellite. As a result, the latter can track the debris, to provide supplementary observations (in addition to those given by ground stations) that allow improving the estimation of its trajectory. Instead, in scenario (b) an impact is predicted, therefore the satellite shall perform a collision avoidance maneuver. This is designed with the objective of minimizing propellant consumption, while ensuring a miss distance greater than a specified threshold value. This research considers the overall dynamics by modeling the spacecraft as a multibody structure, with the use of the Kane's method, which simplifies the process of deriving all the governing equations, while identifying their minimum number. Flexibility is introduced using a modal decomposition technique, under the assumption of small amplitude oscillations. In the two scenarios of interest, efficient strategies are introduced to rotate the solar panels, for the purpose of either maximizing their irradiation or minimizing their oscillations. A new nonlinear reduced-attitude control law, which enjoys global stability properties, is proposed, and is shown to improve the performance of attitude maneuvers when only a single axis must be driven toward a desired direction. Actuation is modeled as well, by assuming the use of a pyramidal array of four single-gimbal control momentum gyroscopes. Their steering law includes singularity avoidance, based on the singular value decomposition of the actuation matrix. Numerical simulations demonstrate that the agile attitude maneuvering strategies proposed in this study allow achieving the operational objectives in the two scenarios of interest, with limited elastic oscillations.

Electrothermally Driven Reconfiguration of Microrobotic Beam Structures for the ChipSail System

Solar sailing enables efficient propellant-free attitude adjustment and orbital maneuvers of solar sail spacecraft with high area-to-mass ratios. However, the heavy supporting mass for large solar sails inevitably leads to low area-to-mass ratios. Inspired by chip-scale satellites, a chip-scale solar sail system named ChipSail, consisting of microrobotic solar sails and a chip-scale satellite, was proposed in this work. The structural design and reconfigurable mechanisms of an electrothermally driven microrobotic solar sail made of Al\Ni50Ti50 bilayer beams were introduced, and the theoretical model of its electro-thermo-mechanical behaviors was established. The analytical solutions to the out-of-plane deformation of the solar sail structure appeared to be in good agreement with the finite element analysis (FEA) results. A representative prototype of such solar sail structures was fabricated on silicon wafers using surface and bulk microfabrication, followed by an in-situ experiment of its reconfigurable property under controlled electrothermal actuation. The experimental results demonstrated significant electro-thermo-mechanical deformation of such microrobotic bilayer solar sails, showing great potential in the development of the ChipSail system. Analytical solutions to the electro-thermo-mechanical model, as well as the fabrication process and characterization techniques, provided a rapid performance evaluation and optimization of such microrobotic bilayer solar sails for the ChipSail.

Multiobjective Design of Gravity-Assist Trajectories via Graph Transcription and Dynamic Programming

Multiple gravity-assist (MGA) trajectory design requires the solution of a mixed-integer programming problem to find the best sequence among all possible combinations of candidate planets and dates for spacecraft maneuvers. Current approaches require computing times rising steeply with the number of control parameters, and they strongly rely on narrow search spaces. Moreover, the challenging multiobjective optimization needs to be tackled to appropriately inform the mission design with full extent of launch opportunities. This paper describes a methodology based upon a trajectory model to transcribe the mixed-integer space into a discrete graph made by grids of interconnected nodes. The model is based on Lambert arc grids obtained for a range of departure dates and flight times between two planets. A Tisserand-based criterion selects planets to pass by. Dynamic programming is extended to multiobjective optimization of MGA trajectories and used to explore the graph, guaranteeing Pareto optimality with only moderate computational effort. Robustness is ensured by evaluating the relationship between graph and mixed-integer spaces. Missions to Jupiter and Saturn alongside challenging comet sample return transfers involving long MGA sequences are discussed. These examples illustrate the robustness and efficiency of the proposed approach in capturing globally optimal solutions and wide Pareto fronts on complex search spaces.

In-flight validation of the Metis visible-light polarimeter coronagraph on board Solar Orbiter

Context. The Metis coronagraph is one of the remote sensing instruments of the ESA-NASA Solar Orbiter mission. The goal for the instrument is to enable the study of the solar atmosphere and solar wind by simultaneously acquiring images of the solar corona at two different wavelengths: visible light (VL), within a band ranging from 580 nm to 640 nm, and ultraviolet light, in the HI Lyα 121.6 ± 10 nm. The visible-light channel of the coronagraph includes a polarimeter with electro-optically modulating liquid crystal variable retarders to measure the linearly polarized brightness of the K-corona and derive the electron density. Aims. In this paper, we present the first in-flight validation results of the Metis polarimetric channel together with a comparison to the on-ground calibrations. This paper seeks to validate the first use of an electro-optical device, the liquid crystal-based polarimeter, in deep space and within a hard radiation environment. Methods. We used the orientation of the K-corona’s linear polarization vector during the roll maneuvers of the Space Orbiter spacecraft for the in-flight calibration. Results. The Metis coronagraph on board the Solar Orbiter shows good agreement with the on-ground measurements. The in-flight validation confirms the expected performance of the visible-light channel polarimeter. Furthermore, a comparison between the first polarized brightness value obtained by Metis and the polarized brightness values obtained by the space-based coronagraph LASCO and the ground-based coronagraph K-Cor shows the consistency of the Metis calibrated results.

An experimental maneuver control for rendezvous and autonomous docking of a small spacecraft

This paper deals with maneuver control for the autonomous docking of two small spacecraft in a rendezvous flight. Due to hardware constraints, maneuver control suppressing fuel consumption and computational cost is a significant issue for small spacecraft. Here, the maneuver control for the approaching motion used a model predictive control system. The spacecraft's maneuvers in two-dimensional plane motion were performed in a frictionless environment with air bearings to verify the control performance.

Learning Reference Governor for Constrained Spacecraft Rendezvous and Proximity Maneuvering

Spacecraft automated rendezvous, proximity maneuvering, and docking (ARPOD) play significant roles in many space missions, including on-orbit servicing and active debris removal. Precise modeling and prediction of spacecraft dynamics can be challenging due to the uncertainties and perturbation forces in the spacecraft operating environment and due to the multilayered structure of its nominal control system. Despite this complication, spacecraft maneuvers need to satisfy required constraints (thrust limits, line-of-sight cone constraints, relative velocity of approach constraints, etc.) to ensure safety and achieve ARPOD objectives. This paper considers an application of a learning-based reference governor (LRG) to spacecraft ARPOD operations to enforce constraints without relying on a dynamic model of the spacecraft during the mission. Similar to the conventional reference governor (RG), the LRG is an add-on supervisor to a closed-loop control system, serving as a prefilter on the command generated by the ARPOD planner. The LRG modifies, if it becomes necessary, the reference command to a constraint-admissible value to enforce specified constraints. The LRG is distinguished, however, by the ability to rely on learning instead of an explicit model of the system; and it guarantees constraints’ satisfaction during and after the learning. In this paper, the LRG is applied to the control of combined translational and rotational motions of a chaser spacecraft, and three case studies with different sets of safety constraints and thruster assumptions are used to demonstrate the benefits of the LRG in ARPOD missions.

Detection and characterisation of unknown maneuvers in spacecraft

Space debris detection and characterisation is an important and pressing field, as the amount of space debris is growing due to the increase in spacecraft launches and increasing interests for space uses. Space agencies like NASA and ESA keep track of over 5,000 spacecraft and 20,000 spacecraft-related debris objects currently orbiting the Earth. An active spacecraft may undergo trajectory changes through an impulsive maneuver, or an inactive spacecraft may change its trajectory because of uncontrolled explosions due to residual fuel without warning. These pose risks to operational satellites. In addition, measurement data from tracking and detection consists of uncertainties, causing maneuver detection and characterisation methods to become computationally expensive.Thus, the goal of this work was to develop a quick, reliable, computationally inexpensive method for spacecraft maneuver detection and characterisation to be used by satellite operators and assist in collision avoidance measures. This paper presents such a technique, which is applied to simulated measurement data of an orbiting spacecraft taken from one ground site. With these objectives, a method using a modified Extended Kalman Filter (EKF) with covariance inflation from previous studies was explored and simplified. A new parameter definition is presented for maneuver detection and characterisation, named the Kaderali-parameter, or K, with focus on transverse apogee and perigee maneuvers. A relationship was then found between the parameter and the magnitude of the impulsive artificial maneuver. The method was tested in several different scenarios, varying Δv magnitude and scale, Δv direction, maneuver orbital location, orbit geometry, noise covariance, detection parameter threshold, state covariance inflation threshold, and maneuver duration.In general, it was found that as the Δv magnitude increases, the magnitude of the detection parameter also increases. This relationship was found to be a nonlinear logarithmic relationship through best fit curve testing. Possible extensions of this work in the future include: testing multi-maneuver scenarios, multi-sensor data collection and fusion, break-up or berthing events, rendezvous or mating events, and matlab EKF protocol implementation.