The Experts below are selected from a list of 7518 Experts worldwide ranked by ideXlab platform
Simone Damico - One of the best experts on this subject based on the ideXlab platform.
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spacecraft swarm dynamics and control about asteroids
Advances in Space Research, 2020Co-Authors: Corinne Lippe, Simone DamicoAbstract:Abstract This paper presents a novel methodology to control spacecraft swarms about single asteroids. This approach enables the use of small, autonomous swarm spacecraft in conjunction with a mothership, reducing the need for the Deep Space Network and improving performance in future asteroid missions. The methodology is informed by a semi-analytical model for the spacecraft relative motion that includes relevant gravitational effects without assuming J2-dominance as well as solar radiation pressure. The dynamics model is exploited in an Extended Kalman Filter (EKF) to produce an osculating-to-mean relative Orbital Element (ROE) conversion that relies on minimum knowledge of the asteroid gravity. The resulting real-time relative mean state estimate is utilized in a new formation-keeping control algorithm. The control problem is cast in mean relative Orbital Elements to leverage the geometric insight of secular and long-period effects in the definition of control windows for swarm maintenance. Analytical constraints that ensure collision avoidance and enforce swarm geometry are derived and enforced in ROE space. The proposed swarm-keeping algorithms are tested and validated in high-fidelity simulations for a reference asteroid mission.
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spacecraft swarm dynamics and control about asteroids
arXiv: Space Physics, 2020Co-Authors: Corinne Lippe, Simone DamicoAbstract:This paper presents a novel methodology to control spacecraft swarms about singleasteroids with arbitrary gravitational potential coefficients. This approach enablesthe use of small, autonomous swarm spacecraft in conjunction with a mothership,reducing the need for the Deep Space Network and increasing safety in future as-teroid missions. The methodology is informed by a semi-analytical model for thespacecraft absolute and relative motion that includes relevant gravitational effectswithout assuming J2-dominance as well as solar radiation pressure. The dynamicsmodel is exploited in an Extended Kalman Filter (EKF) to produce an osculating-to-mean relative Orbital Element (ROE) conversion that is asteroid agnostic. Theresulting real-time relative mean state estimate is utilized in the formation-keepingcontrol algorithm. The control problem is cast in mean relative Orbital Elements toleverage the geometric insight of secular and long-period effects in the definitionof control windows for swarm maintenance. Analytical constraints that ensurecollision avoidance and enforce swarm geometry are derived and enforced in ROEspace. The proposed swarm-keeping algorithms are tested and validated in high-fidelity simulations for a reference asteroid mission.
Corinne Lippe - One of the best experts on this subject based on the ideXlab platform.
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spacecraft swarm dynamics and control about asteroids
Advances in Space Research, 2020Co-Authors: Corinne Lippe, Simone DamicoAbstract:Abstract This paper presents a novel methodology to control spacecraft swarms about single asteroids. This approach enables the use of small, autonomous swarm spacecraft in conjunction with a mothership, reducing the need for the Deep Space Network and improving performance in future asteroid missions. The methodology is informed by a semi-analytical model for the spacecraft relative motion that includes relevant gravitational effects without assuming J2-dominance as well as solar radiation pressure. The dynamics model is exploited in an Extended Kalman Filter (EKF) to produce an osculating-to-mean relative Orbital Element (ROE) conversion that relies on minimum knowledge of the asteroid gravity. The resulting real-time relative mean state estimate is utilized in a new formation-keeping control algorithm. The control problem is cast in mean relative Orbital Elements to leverage the geometric insight of secular and long-period effects in the definition of control windows for swarm maintenance. Analytical constraints that ensure collision avoidance and enforce swarm geometry are derived and enforced in ROE space. The proposed swarm-keeping algorithms are tested and validated in high-fidelity simulations for a reference asteroid mission.
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spacecraft swarm dynamics and control about asteroids
arXiv: Space Physics, 2020Co-Authors: Corinne Lippe, Simone DamicoAbstract:This paper presents a novel methodology to control spacecraft swarms about singleasteroids with arbitrary gravitational potential coefficients. This approach enablesthe use of small, autonomous swarm spacecraft in conjunction with a mothership,reducing the need for the Deep Space Network and increasing safety in future as-teroid missions. The methodology is informed by a semi-analytical model for thespacecraft absolute and relative motion that includes relevant gravitational effectswithout assuming J2-dominance as well as solar radiation pressure. The dynamicsmodel is exploited in an Extended Kalman Filter (EKF) to produce an osculating-to-mean relative Orbital Element (ROE) conversion that is asteroid agnostic. Theresulting real-time relative mean state estimate is utilized in the formation-keepingcontrol algorithm. The control problem is cast in mean relative Orbital Elements toleverage the geometric insight of secular and long-period effects in the definitionof control windows for swarm maintenance. Analytical constraints that ensurecollision avoidance and enforce swarm geometry are derived and enforced in ROEspace. The proposed swarm-keeping algorithms are tested and validated in high-fidelity simulations for a reference asteroid mission.
Pini Gurfil - One of the best experts on this subject based on the ideXlab platform.
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Optimal Geostationary Satellite Collocation Using Relative Orbital Element Corrections
Journal of Spacecraft and Rockets, 2009Co-Authors: Igor Beigelman, Pini GurfilAbstract:The increasing satellite congestion at the geostationary altitude requires positioning a few satellites in the same geostationary slot, a technique known as collocation. In this paper, we develop a geostationary orbit satellite collocation algorithm using relative Orbital Element corrections, which represent the differences between the impulsive Orbital Element corrections of any two spacecraft in a geosynchronous slot. The main idea is that formulating the problem of collocation in terms of relative OrbitalElement corrections leaves some of the final values of the Orbital Elements unconstrained. The freedom rendered by this modeling can be used to find impulsive maneuvers minimizing a given performance index. The minimum distance between satellites that guarantees collision-freemotionisincorporatedintothedesignprocessto findnecessaryandsufficientconditionsfortherelative eccentricity and inclination vectors, guaranteeing safe collocation. The proposed collocation algorithm is illustrated in a simulation.
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Optimal Fuel-Balanced Impulsive Formationkeeping for Perturbed Spacecraft Orbits
Journal of Guidance Control and Dynamics, 2008Co-Authors: Igor Beigelman, Pini GurfilAbstract:This paper develops an impulsive spacecraft formation-flying control algorithm using relative-Orbital-Element corrections. This formalism introduces an inherent freedom that is used for deriving an optimal formationkeeping law, balancing the fuel consumption among the spacecraft based on the impulsive Gauss variational equations. The main idea is that formulating the problem of formationkeeping in terms of relative-Orbital-Element corrections leaves the final values of the Orbital Elements unconstrained, thus allowing the spacecraft to create a natural energy-balanced formation. The freedom rendered by this modeling is used to find optimal impulsive maneuvers, minimizing the squared 12-norm of the velocity-correction vector, which can be used for formation initialization and control. The optimization is solved using the least-squares method. The optimal formationkeeping method is designed to accommodate the effects of oblateness and drag. Based on graph theory, it is shown that the spacecraft will naturally form a stable energy-balanced formation and that the optimal formationkeeping strategy is invariant to the spanning tree. The algorithm is illustrated by simulating the motion of a formation of spacecraft possessing different ballistic coefficients subject to oblateness and drag.
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Manifolds and Metrics in the Relative Spacecraft Motion Problem
Journal of Guidance Control and Dynamics, 2006Co-Authors: Pini Gurfil, Konstantin V. KholshevnikovAbstract:This paper establishes a methodology for obtaining the general solution to the spacecraft relative motion problem by utilizing the Cartesian configuration space in conjunction with classical Orbital Elements. The geometry of the relative motion configuration space is analyzed, and the relative motion invariant manifold is determined. Most importantly, the geometric structure of the relative motion problem is used to derive useful metrics for quantification of the minimum, maximum, and mean distance between spacecraft for commensurable and noncommensurable mean motions. A number of analytic solutions as well as useful examples are provided, illustrating the calculated bounds. A few particular cases that yield simple solutions are given. Nomenclature a = semimajor axis E = eccentric anomaly E = follower orbit e = eccentricity F = follower perifocal frame f = true anomaly I = inertial frame i = inclination Jk = Bessel function L = leader-fixed frame M = mean anomaly n = mean motion n0 = fundamental frequency R = leader position vector R = relative motion invariant manifold r = follower position vector W = distance function α = normalized semimajor axis μ = gravitational constant ρ = relative position vector � = right ascension of the ascending node ω = argument of periapsis ω = angular velocity vector |·| = vector norm �·� = signal norm Superscripts � = leader ∗ = relative Orbital Element
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Accessibility, stabilizability, and feedback control of continuous Orbital transfer.
Annals of the New York Academy of Sciences, 2004Co-Authors: Pini GurfilAbstract:: This paper investigates the problem of low-thrust Orbital transfer using Orbital Element feedback from a control-theoretic standpoint, concepts of controllability, feedback stabilizability, and their interaction. The Gauss variational equations (GVEs) are used to model the state-space dynamics. First, the notion of accessibility, a weaker form of controllability, is presented. It is then shown that the GVEs are globally accessible. Based on the accessibility result, a nonlinear feedback controller is derived that asymptotically steers a vehicle from an initial elliptic Keplerian orbit to any given elliptic Keplerian orbit. The performance of the new controller is illustrated by simulating an Orbital transfer between two geosynchronous Earth orbits. It is shown that the low-thrust controller requires less fuel than an impulsive maneuver for the same transfer time. Closed-form, analytic expressions for the new Orbital transfer controller are given. Finally, it is proved, based on a topological nonlinear stabilizability test, that there does not exist a continuous closed-loop controller that can transfer a spacecraft to a parabolic escape trajectory.
A. Vienne - One of the best experts on this subject based on the ideXlab platform.
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Astronomy Astrophysics Analytical representation for ephemeride with short time spans
Astronomy and Astrophysics - A&A, 2020Co-Authors: A. VienneAbstract:Context. The ephemerides of natural satellites resulting from numerical integration have a very good precision on the fitting to recent observations, in a limited interval. Meanwhile, synthetic ephemerides like the Théorie Analytique des Satellites de Saturne (TASS) by Vienne and Duriez describe in detail the dynamical system by a representation based on the combinations of the proper frequencies. Some theoretical studies need to have both advantages. For example, to study the rotation of Titan, one needs to know the representation of its longitude. Aims. We aim to use these two types of ephemerides in order to rebuild a long-lasting and high-precision ephemeris with proper frequencies based on the numerical integration ephemeris. The aim is to describe the numerical ephemerides with formulas similar to analytical ones. Methods. We used the representation of the Orbital Elements from the TASS ephemeris analysed over 10 000 years as a reference template. We obtained the proper frequencies with both numerical and the TASS ephemeris over 1000 years only. A least-square procedure allowed us to get the analytical representation of an Orbital Element in this limited interval. Results. We acquire the representation of the mean longitude of Titan from JPL ephemeris over 1000 years. For almost all components, the corresponding amplitudes and phases are similar to the relative terms from TASS. The biggest difference between our representation and the mean longitude of Titan of JPL is less than 100 km over 1000 years, and the standard deviation is about 26 km.
Colin R. Mcinnes - One of the best experts on this subject based on the ideXlab platform.
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Secular Orbital Element Variations Due to Continuous Low-thrust Control and Third-body Perturbations
2019Co-Authors: Xiaoyu Liu, Colin R. Mcinnes, Matteo CeriottiAbstract:An analytical method is investigated to evaluate the low-thrust orbit dynamics of a spacecraft under the influence of third-body perturbations in the absence of resonance. The Orbital evolution is formulated as a combination of Lagragian and Gaussian variational equations with two distinct perturbing accelerations. The third-body effects are described by Legendre polynomials and truncated up to second-order terms. Components of the control terms are represented as Fourier series in eccentric anomaly. Then the variational equations are averaged over its period to yield secular effects. The work is also presented in non-singular Elements. Numerical simulation verifies the efficiency of the methodology.
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Asteroid Resource Map for Near-Earth Space
Journal of Spacecraft and Rockets, 2011Co-Authors: Joan-pau Sanchez Cuartielles, Colin R. McinnesAbstract:Most future concepts for the exploration and exploitation of space require a large initial mass in low Earth orbit. Delivering this required mass from the Earth’s surface increases cost due to the large energy input necessary to move mass out of the Earth’s gravity well. An alternative is to search for resources in-situ among the near Earth asteroid population. The near Earth asteroid resources that could be transferred to a bound Earth orbit are determined by integrating the probability of finding asteroids inside the Keplerian Orbital Element space of the set of transfers with an specific energy smaller than a given threshold. Transfers are defined by a series of impulsive maneuvers and computed using the patched-conic approximation. The results show that even moderately low energy transfers enable access to a large mass of resources.
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Accessibility of the resources of near Earth space using multi-impulse transfers
AIAA AAS Astrodynamics Specialist Conference, 2010Co-Authors: J.p. Sanchez, Colin R. McinnesAbstract:Most future concepts for exploration and exploitation of space require a large initial mass in low Earth orbit. Delivering this mass requires overcoming Earth's natural gravity well, which imposes a distinct obstacle to space-faring. An alternative for future space progress is to search for resources in-situ among the near Earth asteroid population. This paper examines the scenario of future utilization of asteroid resources. The near Earth asteroid resources that could be transferred to a bound Earth orbit are determined by integrating the probability of finding asteroids inside the Keplerian Orbital Element space of the set of transfers with an specific energy smaller than a given threshold. Transfers are defined by a series of impulsive maneuvers and computed using the patched-conic approximation. The results show that even moderately low energy transfers enable access to a large mass of resources.