Analysis of the relative dynamics of a charged spacecraft moving under the influence of a magnetic field

We analyse a dynamical scenario where a constantly charged spacecraft (follower) moves in the vicinity of another one (leader) that follows a circular Keplerian orbit around the Earth and generates a rotating magnetic dipole. The mass of the follower is assumed to be negligible when compared with the one of the leader and they are supposed to be in a high-Earth orbit, so the Lorentz force on the follower due to the geomagnetic field is ignored. With these assumptions, the motion of the leader is not perturbed by the follower and it is only subjected to the Earth’s gravitational force field, while the charged follower is under to both the gravitational force of the Earth and the Lorentz force due to the magnetic dipole of the leader. We focus on the dynamical characteristics of the system as a function of its parameters, with special attention to the ratio of the leader’s mean motion around the Earth to the rotating rate of the dipole. We study the critical points of the model and their stability, the admissible and forbidden regions of motion of the deputy using the zero velocity surfaces and the families of periodic orbits emanating from equilibria. In the normal case we pay special attention to the periodic orbits of elliptic type and to the families of 2D tori surrounding them that are computed by means of a parameterisation method. The result is a fine catalog of orbits together with an accurate dynamical description suitable to researchers interested in potential applications of satellite formation flight using this kind of technology.Peer ReviewedPostprint (author's final draft

Optimal Guidance and Control for Electromagnetic Formation Flying

学位の種別: 修士University of Tokyo(東京大学

TeSeR – Technology for Self-Removal – Status of a Horizon 2020 project to ensure the Post-Mission-Disposal of any future spacecraft

No abstract available

2020 NASA Technology Taxonomy

This document is an update (new photos used) of the PDF version of the 2020 NASA Technology Taxonomy that will be available to download on the OCT Public Website. The updated 2020 NASA Technology Taxonomy, or "technology dictionary", uses a technology discipline based approach that realigns like-technologies independent of their application within the NASA mission portfolio. This tool is meant to serve as a common technology discipline-based communication tool across the agency and with its partners in other government agencies, academia, industry, and across the world

Optimal Finite Thrust Guidance Methods for Constrained Satellite Proximity Operations Inspection Maneuvers

Algorithms are developed to find optimal guidance for an inspector satellite operating nearby a resident space object (RSO). For a non-maneuvering RSO, methods are first developed for a satellite subject to maximum slew rates to conduct an initial inspection of an RSO, where the control variables include the throttle level and direction of the thrust. Second, methods are developed to optimally maneuver a satellite with on/off thrusters into a natural motion circumnavigation or teardrop trajectory, subject to lighting and collision constraints. It is shown that for on/off thrusters, a control sequence can be parameterized to a relatively small amount of control variables and the relative states can be analytically propagated as a function of those control variables. For a maneuvering RSO, differential games are formulated and solved for an inspector satellite to achieve multiple inspection goals, such as aligning with the Sun vector or matching the RSO\u27s energy. The developed algorithms lead to fuel and time savings which can increase the mission life and capabilities of inspector satellites and thus improve space situational awareness for the U.S. Air Force

Attitude and formation control design and system simulation for a three-satellite CubeSat mission

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2013.This electronic version was submitted and approved by the author's academic department as part of an electronic thesis pilot project. The certified thesis is available in the Institute Archives and Special Collections.Cataloged from department-submitted PDF version of thesisIncludes bibliographical references (p. 113-115).Spacecraft formation flight has been identified as a critical enabling technology for achieving many scientific, commercial, and military objectives. One of the primary challenges of a formation flight mission is the control of the relative motion between spacecraft. Before any flagship missions will launch, technology development missions will be required to demonstrate the utility and functionality of formation flying systems. This thesis describes the complete attitude and formation control design for the MotherCube formation flight technology demonstration mission in LEO. A model of the spacecraft's sensors and actuators is developed and analyzed. Using curvilinear orbit theory, a simple LQR control law is used to generate a set of desired relative accelerations for formation control. A newly developed two-tier numerical allocation scheme is used alongside an independent PD attitude control law to generate a set of actuator commands which provides 3-axis attitude stabilization as well as formation control with guaranteed feasibility of actuator commands. An Extended Kalman Filter was developed to estimate the system attitude and angular rate from sensor measurements. To test these algorithms, a simulation environment was developed. This environment includes realistic models of space environment and the major perturbation effects which a LEO spacecraft formation would encounter. In order to improve the fidelity, a new intermediate-accuracy method for computing attitude-dependent aerodynamic and solar effects was also developed. Finally, results from the simulation are used numerically validate the dual-allocator approach, assess the performance of the control laws and provide system level metrics such as fuel use and required maneuver time.by Austin Kyle Nicholas.S.M

Autonomous formation flying: unified control and collision avoidance methods for close manoeuvring spacecraft

The idea of spacecraft formations, flying in tight configurations with maximum baselines of a few hundred meters in low-Earth orbits, has generated widespread interest over the last several years. Nevertheless, controlling the movement of spacecraft in formation poses difficulties, such as in-orbit high-computing demand and collision avoidance capabilities, which escalate as the number of units in the formation is increased and complicated nonlinear effects are imposed to the dynamics, together with uncertainty which may arise from the lack of knowledge of system parameters. These requirements have led to the need of reliable linear and nonlinear controllers in terms of relative and absolute dynamics. The objective of this thesis is, therefore, to introduce new control methods to allow spacecraft in formation, with circular/elliptical reference orbits, to efficiently execute safe autonomous manoeuvres. These controllers distinguish from the bulk of literature in that they merge guidance laws never applied before to spacecraft formation flying and collision avoidance capacities into a single control strategy. For this purpose, three control schemes are presented: linear optimal regulation, linear optimal estimation and adaptive nonlinear control. In general terms, the proposed control approaches command the dynamical performance of one or several followers with respect to a leader to asymptotically track a time-varying nominal trajectory (TVNT), while the threat of collision between the followers is reduced by repelling accelerations obtained from the collision avoidance scheme during the periods of closest proximity. Linear optimal regulation is achieved through a Riccati-based tracking controller. Within this control strategy, the controller provides guidance and tracking toward a desired TVNT, optimizing fuel consumption by Riccati procedure using a non-infinite cost function defined in terms of the desired TVNT, while repelling accelerations generated from the CAS will ensure evasive actions between the elements of the formation. The relative dynamics model, suitable for circular and eccentric low-Earth reference orbits, is based on the Tschauner and Hempel equations, and includes a control input and a nonlinear term corresponding to the CAS repelling accelerations. Linear optimal estimation is built on the forward-in-time separation principle. This controller encompasses two stages: regulation and estimation. The first stage requires the design of a full state feedback controller using the state vector reconstructed by means of the estimator. The second stage requires the design of an additional dynamical system, the estimator, to obtain the states which cannot be measured in order to approximately reconstruct the full state vector. Then, the separation principle states that an observer built for a known input can also be used to estimate the state of the system and to generate the control input. This allows the design of the observer and the feedback independently, by exploiting the advantages of linear quadratic regulator theory, in order to estimate the states of a dynamical system with model and sensor uncertainty. The relative dynamics is described with the linear system used in the previous controller, with a control input and nonlinearities entering via the repelling accelerations from the CAS during collision avoidance events. Moreover, sensor uncertainty is added to the control process by considering carrier-phase differential GPS (CDGPS) velocity measurement error. An adaptive control law capable of delivering superior closed-loop performance when compared to the certainty-equivalence (CE) adaptive controllers is finally presented. A novel noncertainty-equivalence controller based on the Immersion and Invariance paradigm for close-manoeuvring spacecraft formation flying in both circular and elliptical low-Earth reference orbits is introduced. The proposed control scheme achieves stabilization by immersing the plant dynamics into a target dynamical system (or manifold) that captures the desired dynamical behaviour. They key feature of this methodology is the addition of a new term to the classical certainty-equivalence control approach that, in conjunction with the parameter update law, is designed to achieve adaptive stabilization. This parameter has the ultimate task of shaping the manifold into which the adaptive system is immersed. The performance of the controller is proven stable via a Lyapunov-based analysis and Barbalat’s lemma. In order to evaluate the design of the controllers, test cases based on the physical and orbital features of the Prototype Research Instruments and Space Mission Technology Advancement (PRISMA) are implemented, extending the number of elements in the formation into scenarios with reconfigurations and on-orbit position switching in elliptical low-Earth reference orbits. An extensive analysis and comparison of the performance of the controllers in terms of total Δv and fuel consumption, with and without the effects of the CAS, is presented. These results show that the three proposed controllers allow the followers to asymptotically track the desired nominal trajectory and, additionally, those simulations including CAS show an effective decrease of collision risk during the performance of the manoeuvre

The dynamics and control of large space structures with distributed actuation

Future large space structures are likely to be constructed at much greater length-scales, and lower areal mass densities than has been achieved to-date. This could be enabled by ongoing developments in on-orbit manufacturing, whereby large structures are 3D-printed in space from raw feedstock materials. This thesis proposes and analyses a number of attitude control strategies which could be adopted for this next generation of ultra-lightweight, large space structures. Each of the strategies proposed makes use of distributed actuation, which is demonstrated early in the thesis to reduce structural deformations during attitude manoeuvres. All of the proposed strategies are considered to be particularly suitable for structures which are 3d-printed on-orbit, due to the relative simplicity of the actuators and ease with which the actuator placement or construction could be integrated with the on-orbit fabrication of the structure itself. The first strategy proposed is the use of distributed arrays of magnetorquer rods. First, distributed torques are shown to effectively rotate highly flexible structures. This is compared with torques applied to the centre-of-mass of the structure, which cause large surface deformations and can fail to enact a rotation. This is demonstrated using a spring-mass model of a planar structure with embedded actuators. A torque distribution algorithm is then developed to control an individually addressable array of actuators. Attitude control simulations are performed, using the array to control a large space structure, again modelled as a spring-mass system. The attitude control system is demonstrated to effectively detumble a representative 75×75m flexible structure, and perform slew manoeuvres, in the presence of both gravity-gradient torques and a realistic magnetic field model. The development of a Distributed Magnetorquer Demonstration Platform is then presented, a laboratory-scale implementation of the distributed magnetorquer array concept. The platform consists of 48 addressable magnetorquers, arranged with two perpendicular torquers at the nodes of a 5×5 grid. The control algorithms proposed previously in the thesis are implemented and tested on this hardware, demonstrating the practical feasibility of the concept. Results of experiments using a spherical air bearing and Helmholtz cage are presented, demonstrating rest-to-rest slew manoeuvres and detumbling around a single axis using the developed algorithms. The next attitude control strategy presented is the use of embedded current loops, conductive pathways which can be integrated with a spacecraft support structure and used to generate control torques through interaction with the Earth’s magnetic field. Length-scaling laws are derived by determining what fraction of a planar spacecraft’s mass would need to be allocated to the conductive current loops in order to produce a torque at least as large as the gravity gradient torque. Simulations are then performed of a flexible truss structure, modelled as a spring-mass system, for a range of structural flexibilities and a variety of current loop geometries. Simulations demonstrate rotation of the structure via the electromagnetic force on the current carrying elements, and are also used to characterise the structural deformations caused by the various current loop geometries. An attitude control simulation is then performed, demonstrating a 90◦ slew manoeuvre of a 250×250 m flexible structure through the use of three orthogonal sets of current loops embedded within the spacecraft. The final concept investigated in this thesis is a self-reconfiguring OrigamiSat, where reconfiguration of the proposed OrigamiSat is triggered by changes in the local surface optical properties of an origami structure to harness the solar radiation pressure induced acceleration. OrigamiSats are origami spacecraft with reflective panels which, when flat, operate as a conventional solar sail. Shape reconfiguration, i.e. “folding” of the origami design, allows the OrigamiSat to change operational modes, performing different functions as per mission requirements. For example, a flat OrigamiSat could be reconfigured into the shape of a parabolic reflector, before returning to the flat configuration when required to again operate as a solar sail, providing propellant-free propulsion. Shape reconfiguration or folding of OrigamiSats through the use of surface reflectivity modulation is investigated in this thesis. First, a simplified, folding facet model is used to perform a length-scaling analysis, and then a 2d multibody dynamics simulation is used to demonstrate the principle of solar radiation presure induced folding. A 3d multibody dynamics simulation is then developed and used to demonstrate shape reconfiguration for different origami folding patterns. Here, the attitude dynamics and shape reconfiguration of OrigamiSats are found to be highly coupled, and thus present a challenge from a control perspective. The problem of integrating attitude and shape control of a Miura-fold pattern OrigamiSat through the use of variable reflectivity is then investigated, and a control algorithm developed which uses surface reflectivity modulation of the OrigamiSat facets to enact shape reconfiguration and attitude manoeuvres simultaneously