The Experts below are selected from a list of 240 Experts worldwide ranked by ideXlab platform
Daniel J Scheeres - One of the best experts on this subject based on the ideXlab platform.
-
on the a and g families of orbits in the hill problem with solar radiation pressure and their application to asteroids
AIAA AAS Astrodynamics Specialist Conference 2014, 2015Co-Authors: Daniel Garcia Yarnoz, Daniel J Scheeres, Colin R McinnesAbstract:The focus of this paper is on the exploration of the a and g-g’ families of planar symmetric periodic orbits around minor bodies under the effect of solar radiation pressure. An extended Hill problem with solar radiation pressure (SRP) allows the study of Spacecraft Trajectories in the vicinity of asteroids orbiting the Sun. The evolution of the a and g-g’ families is presented with SRP increasing from the classical Hill problem to levels characteristic of current and future planned missions to minor bodies, as well as one extreme case with very large SRP for a small asteroid. In addition, the implications of considering a spherical body are analysed, in terms of Trajectories colliding with the asteroid and eclipses, which limits the feasibility of various family branches. Finally, the influence of SRP on the linear stability of feasible orbits is calculated.
-
on the a and g families of orbits in the hill problem with solar radiation pressure and their application to asteroid orbiters
Celestial Mechanics and Dynamical Astronomy, 2015Co-Authors: Daniel Garcia Yarnoz, Daniel J Scheeres, Colin R McinnesAbstract:The focus of this paper is on the exploration of the \(a\) and \(g\)-\(g'\) families of planar symmetric periodic orbits around minor bodies under the effect of solar radiation pressure (SRP). An extended Hill problem with SRP allows the study of Spacecraft Trajectories in the vicinity of asteroids orbiting the Sun. The evolution of the \(a\) and \(g\)-\(g'\) families is presented with SRP increasing from the classical Hill problem to levels characteristic of current and future planned missions to minor bodies, as well as one extreme case with very large SRP for a small asteroid. In addition, the implications of considering a spherical body are analysed, in terms of Trajectories colliding with the asteroid and eclipses, which limits the feasibility of various family branches. Finally, the influence of SRP on the linear stability of feasible orbits is calculated.
-
abstraction predictive control for chaotic Spacecraft orbit design
IFAC-PapersOnLine, 2015Co-Authors: David Allen Surovik, Daniel J ScheeresAbstract:Abstract We formulate a predictive control algorithm for planning Spacecraft Trajectories at close proximity to asteroids, where state-space dynamics are sensitive, complex, and non-periodic, thus invalidating standard orbit design methods. Sets of scientific observation objectives are pursued by defining an abstracted goal space and a control law for regulating the motion in this space toward a state signifying mission completion. Non-deterministic components of the control scheme, which obtains suboptimal solutions via heuristic search of the complex control space, are found to produce diverse solution paths through the state space and goal space alike, implying high sensitivity of the approach despite its ultimate stability. The degree of this diversity is charted as a function of chaoticity in the state-space dynamics and the complexity of the goal-space definition. These results are compared to those obtained under Keplerian motion and are discussed in the context of eventual application of the algorithm for preliminary mission design and onboard implementation for robust online planning.
-
multiple gravity assists capture and escape in the restricted three body problem
Siam Journal on Applied Dynamical Systems, 2007Co-Authors: Shane D Ross, Daniel J ScheeresAbstract:For low energy Spacecraft Trajectories such as multimoon orbiters for the Jupiter system, multiple gravity assists by moons could be used in conjunction with ballistic capture to drastically decrease fuel usage. In this paper, we investigate a special class of multiple gravity assists which can occur outside of the perturbing body's sphere of influence (the Hill sphere) and which is dynamically connected to orbits that get captured by the perturber and orbits which escape to infinity. We proceed by deriving a family of symplectic twist maps to approximate a particle's motion in the planar circular restricted three-body problem. The maps capture well the dynamics of the full equations of motion; the phase space contains a connected chaotic zone where intersections between unstable resonant orbit manifolds provide the template for lanes of fast migration between orbits of different semimajor axes. Within the chaotic zone, the concept of a set of reachable orbits is useful. This set can be considered bounded...
-
deflection of Spacecraft Trajectories as a new test of general relativity determining the parametrized post newtonian parameters β and γ
Physical Review D, 2004Co-Authors: James M Longuski, Daniel J Scheeres, Ephraim Fischbach, Giacomo Giampieri, R S ParkAbstract:In a previous work, we proposed a new test of general relativity ~GR! based on a general deflection formula which applies to all values of asymptotic speed V‘ (0
William J Mason - One of the best experts on this subject based on the ideXlab platform.
-
optimal multi objective low thrust Spacecraft Trajectories
Computer Methods in Applied Mechanics and Engineering, 2000Co-Authors: Victoria Coverstonecarroll, John W. Hartmann, William J MasonAbstract:Abstract Genetic algorithms have gained popularity as effective search procedures for obtaining solutions to traditionally difficult space mission optimization problems. In this paper, a hybrid optimization method is described that integrates a multi-objective genetic algorithm with a calculus-of-variations-based low-thrust trajectory optimizer. Fronts of Pareto optimal Trajectories are generated and novel Trajectories identified for both Earth–Mars and Earth–Mercury missions.
Colin R Mcinnes - One of the best experts on this subject based on the ideXlab platform.
-
on the a and g families of orbits in the hill problem with solar radiation pressure and their application to asteroids
AIAA AAS Astrodynamics Specialist Conference 2014, 2015Co-Authors: Daniel Garcia Yarnoz, Daniel J Scheeres, Colin R McinnesAbstract:The focus of this paper is on the exploration of the a and g-g’ families of planar symmetric periodic orbits around minor bodies under the effect of solar radiation pressure. An extended Hill problem with solar radiation pressure (SRP) allows the study of Spacecraft Trajectories in the vicinity of asteroids orbiting the Sun. The evolution of the a and g-g’ families is presented with SRP increasing from the classical Hill problem to levels characteristic of current and future planned missions to minor bodies, as well as one extreme case with very large SRP for a small asteroid. In addition, the implications of considering a spherical body are analysed, in terms of Trajectories colliding with the asteroid and eclipses, which limits the feasibility of various family branches. Finally, the influence of SRP on the linear stability of feasible orbits is calculated.
-
on the a and g families of orbits in the hill problem with solar radiation pressure and their application to asteroid orbiters
Celestial Mechanics and Dynamical Astronomy, 2015Co-Authors: Daniel Garcia Yarnoz, Daniel J Scheeres, Colin R McinnesAbstract:The focus of this paper is on the exploration of the \(a\) and \(g\)-\(g'\) families of planar symmetric periodic orbits around minor bodies under the effect of solar radiation pressure (SRP). An extended Hill problem with SRP allows the study of Spacecraft Trajectories in the vicinity of asteroids orbiting the Sun. The evolution of the \(a\) and \(g\)-\(g'\) families is presented with SRP increasing from the classical Hill problem to levels characteristic of current and future planned missions to minor bodies, as well as one extreme case with very large SRP for a small asteroid. In addition, the implications of considering a spherical body are analysed, in terms of Trajectories colliding with the asteroid and eclipses, which limits the feasibility of various family branches. Finally, the influence of SRP on the linear stability of feasible orbits is calculated.
-
designing displaced lunar orbits using low thrust propulsion
Journal of Guidance Control and Dynamics, 2010Co-Authors: Jules Simo, Colin R McinnesAbstract:The design of Spacecraft Trajectories is a crucial task in space mission design. Solar sail technology appears as a promising form of advanced Spacecraft propulsion which can enable exciting new space science mission concepts such as solar system exploration and deep space observation. Although solar sailing has been considered as a practical means of Spacecraft propulsion only relatively recently, the fundamental ideas are by no means new (see McInnes1 for a detailed description). A solar sail is propelled by re ecting solar photons and therefore can transform the momentum of the photons into a propulsive force. This article focuses on designing displaced lunar orbits using low-thrust propulsion.
-
asymptotic analysis of displaced lunar orbits
Journal of Guidance Control and Dynamics, 2009Co-Authors: Jules Simo, Colin R McinnesAbstract:The design of Spacecraft Trajectories is a crucial task in space mission design. Solar sail technology appears as a promising form of advanced Spacecraft propulsion which can enable exciting new space science mission concepts such as solar system exploration and deep space observation. Although solar sailing has been considered as a practical means of Spacecraft propulsion only relatively recently, the fundamental ideas are by no means new (see McInnes1 for a detailed description). A solar sail is propelled by reflecting solar photons and therefore can transform the momentum of the photons into a propulsive force. Solar sails can also be utilised for highly non-Keplerian orbits, such as orbits displaced high above the ecliptic plane (see Waters and McInnes2). Solar sails are especially suited for such non-Keplerian orbits, since they can apply a propulsive force continuously. In such Trajectories, a sail can be used as a communication satellite for high latitudes. For example, the orbital plane of the sail can be displaced above the orbital plane of the Earth, so that the sail can stay fixed above the Earth at some distance, if the orbital periods are equal (see Forward3). Orbits around the collinear points of the Earth-Moon system are also of great interest because their unique positions are advantageous for several important applications in space mission design (see e.g. Szebehely4, Roy,5 Vonbun,6 Thurman et al.,7 Gomez et al.8, 9). Several authors have tried to determine more accurate approximations (quasi-Halo orbits) of such equilibrium orbits10. These orbits were first studied by Farquhar11, Farquhar and Kamel10, Breakwell and Brown12, Richardson13, Howell14, 15.If an orbit maintains visibility from Earth, a Spacecraft on it (near the L2 point) can be used to provide communications between the equatorial regions of the Earth and the lunar poles. The establishment of a bridge for radio communications is crucial for forthcoming space missions, which plan to use the lunar poles.McInnes16 investigated a new family of displaced solar sail orbits near the Earth-Moon libration points.Displaced orbits have more recently been developed by Ozimek et al.17 using collocation methods. In Baoyin and McInnes18, 19, 20 and McInnes16, 21, the authors describe new orbits which are associated with artificial Lagrange points in the Earth-Sun system. These artificial equilibria have potential applications for future space physics and Earth observation missions. In McInnes and Simmons22, the authors investigate large new families of solar sail orbits, such as Sun-centered halo-type Trajectories, with the sail executing a circular orbit of a chosen period above the ecliptic plane. We have recently investigated displaced periodic orbits at linear order in the Earth-Moon restricted three-body system, where the third massless body is a solar sail (see Simo and McInnes23). These highly non-Keplerian orbits are achieved using an extremely small sail acceleration. It was found that for a given displacement distance above/below the Earth-Moon plane it is easier by a factor of order 3.19 to do so at L4=L5 compared to L1=L2 - ie. for a fixed sail acceleration the displacement distance at L4=L5 is greater than that at L1=L2. In addition, displaced L4=L5 orbits are passively stable, making them more forgiving to sail pointing errors than highly unstable orbits at L1=L2.The drawback of the new family of orbits is the increased telecommunications path-length, particularly the Moon-L4 distance compared to the Moon-L2 distance.
Cesar A Ocampo - One of the best experts on this subject based on the ideXlab platform.
-
theoretical foundation of copernicus a unified system for trajectory design and optimization
2010Co-Authors: Cesar A Ocampo, Juan Senent, Jacob WilliamsAbstract:The fundamental methods are described for the general Spacecraft trajectory design and optimization software system called Copernicus. The methods rely on a unified framework that is used to model, design, and optimize Spacecraft Trajectories that may operate in complex gravitational force fields, use multiple propulsion systems, and involve multiple Spacecraft. The trajectory model, with its associated equations of motion and maneuver models, are discussed.
-
finite burn maneuver modeling for a generalized Spacecraft trajectory design and optimization system
Annals of the New York Academy of Sciences, 2004Co-Authors: Cesar A OcampoAbstract:: The modeling, design, and optimization of finite burn maneuvers for a generalized trajectory design and optimization system is presented. A generalized trajectory design and optimization system is a system that uses a single unified framework that facilitates the modeling and optimization of complex Spacecraft Trajectories that may operate in complex gravitational force fields, use multiple propulsion systems, and involve multiple Spacecraft. The modeling and optimization issues associated with the use of controlled engine burn maneuvers of finite thrust magnitude and duration are presented in the context of designing and optimizing a wide class of finite thrust Trajectories. Optimal control theory is used examine the optimization of these maneuvers in arbitrary force fields that are generally position, velocity, mass, and are time dependent. The associated numerical methods used to obtain these solutions involve either, the solution to a system of nonlinear equations, an explicit parameter optimization method, or a hybrid parameter optimization that combines certain aspects of both. The theoretical and numerical methods presented here have been implemented in copernicus, a prototype trajectory design and optimization system under development at the University of Texas at Austin.
David W. Hinckley - One of the best experts on this subject based on the ideXlab platform.
-
evolutionary approach to lambert s problem for non keplerian Spacecraft Trajectories
Aerospace, 2017Co-Authors: David W. Hinckley, Darren L. HittAbstract:In this paper, we use differential evolution (DE), with best-evolved results refined using a Nelder–Mead optimization, to solve boundary-value complex problems in orbital mechanics relevant to low Earth orbits (LEO). A class of Lambert-type problems is examined to evaluate the performance of this evolutionary method in its application to solving nonlinear boundary value problems (BVP) arising in mission planning. In this method, we evolve impulsive initial velocity vectors giving rise to intercept Trajectories that take a Spacecraft from given initial position in space to specified target position. The positional error of the final position is minimized subject to time-of-flight and/or energy (fuel) constraints. The method is first validated by demonstrating its ability to recover known analytical solutions obtainable with the assumption of Keplerian motion; the method is then applied to more complex non-Keplerian problems incorporating trajectory perturbations arising in low Earth orbit (LEO) due to the Earth’s oblateness and rarefied atmospheric drag. The viable Trajectories obtained for these challenging problems demonstrate the ability of this computational approach to handle Lambert-type problems with arbitrary perturbations, such as those occurring in realistic mission trajectory design.
-
multi objective optimization of Spacecraft Trajectories for small body coverage missions
AAS AIAA Space Flight Mechanics Meeting, 2017Co-Authors: David W. Hinckley, Jacob A Englander, Darren L. HittAbstract:Visual coverage of surface elements of a small-body object requires multiple images to be taken that meet many requirements on their viewing angles, illumination angles, times of day, and combinations thereof. Designing Trajectories capable of maximizing total possible coverage may not be useful since the image target sequence and the feasibility of said sequence given the rotation-rate limitations of the Spacecraft are not taken into account. This work presents a means of optimizing, in a multi-objective manner, surface target sequences that account for such limitations.
-
Evolved Non-Keplerian Spacecraft Trajectories for Near-Earth Orbital Maneuvers
AIAA AAS Astrodynamics Specialist Conference, 2014Co-Authors: David W. Hinckley, Darren L. Hitt, Margaret J. EppsteinAbstract:In this paper we use Differential Evolution (DE), with best-evolved results refined using a Nelder-Mead optimization, to solve complex problems in orbital mechanics relevant to low Earth orbits (LEO) and within the Earth-Moon system. A class of Lambert problems is examined to evaluate the performance and robustness of this evolutionary approach to orbit optimization. We evolve impulsive initial velocity vectors giving rise to intercept Trajectories that take a Spacecraft from given initial positions to specified target positions. We seek to minimize final positional error subject to time-of-flight and/or energy (fuel) constraints. We first validate that the method can recover known analytical solutions obtainable with the assumption of Keplerian motion. We then apply the method to more complex and realistic non-Keplerian problems incorporating trajectory perturbations arising in LEO due to the Earth’s oblateness and rarefied atmospheric drag. Finally, a rendezvous trajectory from LEO to the L4 Lagrange point is computed. The viable Trajectories obtained for these challenging problems suggest the robustness of our computational approach for real-world orbital trajectory design in LEO situations where no analytical solution exists.
-
evolved Spacecraft Trajectories for low earth orbit
Genetic and Evolutionary Computation Conference, 2014Co-Authors: David W. Hinckley, Darren L. Hitt, Karol Zieba, Margaret J. EppsteinAbstract:In this paper we use Differential Evolution (DE), with best evolved results refined using a Nelder-Mead optimization, to solve complex problems in orbital mechanics relevant to low Earth orbits (LEO). A class of so-called 'Lambert Problems' is examined. We evolve impulsive initial velocity vectors giving rise to intercept Trajectories that take a Spacecraft from given initial positions to specified target positions. We seek to minimize final positional error subject to time-of-flight and/or energy (fuel) constraints. We first validate that the method can recover known analytical solutions obtainable with the assumption of Keplerian motion. We then apply the method to more complex and realistic non-Keplerian problems incorporating trajectory perturbations arising in LEO due to the Earth's oblateness and rarefied atmospheric drag. The viable Trajectories obtained for these difficult problems suggest the robustness of our computational approach for real-world orbital trajectory design in LEO situations where no analytical solution exists.