Search for a grand tour of the jupiter galilean moons

Izzo, Dario; Simões, Luís; Märtens, Marcus; Croon, Guido C. H. E. de; Héritier, Aurélie; Yam, Chit Hong

doi:10.1145/2463372.2463524

Cited by 28 publications

(17 citation statements)

References 13 publications

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“…The unit direction of this velocity vector is orthogonal to both the periapsis position and the angular momentum, which are both unit vectors: v p =ĥ ×r p (19) The magnitude of the periapsis position and velocity are found next, using another representation for eccentricity and conservation of energy during the flyby:…”

Section: Finding the Close Approach Flyby State At An Intermediate Enmentioning

confidence: 99%

“…Many authors have addressed ballistic patched conics tour design [9,13,18,19,32,33], but few have presented systematic algorithms for representing the results in nbody ephemeris models. Typically, the transition to the full ephemeris model is made in a single leap, and can stress even the best optimization algorithms tackling the simplest of problems.…”

Section: Introductionmentioning

confidence: 99%

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A Continuation Method for Converting Trajectories from Patched Conics to Full Gravity Models

Bradley

Russell

2014

J of Astronaut Sci

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A method is introduced to transition space trajectories from low fidelity patched conics models to full-ephemeris n-body dynamics. The algorithm incorporates a continuation method that progressively re-converges solution trajectories in systems with incremental changes in the dynamics. Continuation is accomplished through the variation of a control parameter, which is tied to body ephemeris locations and masses, minimum flyby altitudes, and sphere of influence sizes. The intermediate models provide a continuous and differentiable path between solutions in the simplified and n-body dynamics, and the boundary values of the control parameter replicate the patched conics and full-ephemeris models exactly. Each successive step preserves the qualitative properties of the initial guess by successively altering flyby states and body masses. A similar approach to this methodology may be taken with any simplified starting guess, such as a restricted three-body model. Trajectories computed using the patched conics conversion method presented here may include gravity assists and rendezvous with any number of target bodies, so the method is ideal for constructing interplanetary or intermoon tour missions, and is amenable to patching shorter segments together to form longer tours.

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Section: Finding the Close Approach Flyby State At An Intermediate Enmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

A Continuation Method for Converting Trajectories from Patched Conics to Full Gravity Models

Bradley

Russell

2014

J of Astronaut Sci

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“…Evolutionary methods utilized have included genetic algorithms [2,3], particle swarm optimization [4], monotonic basin hopping [5], simulated annealing [6], and differential evolution [7]. To a lesser extent, EC methods have also been leveraged for in-orbit trajectory planning about planets [8][9][10].…”

Section: Introductionmentioning

confidence: 99%

“…As such, their approach was necessarily limited to the idealized two-body problem of Keplerian theory and lacked the ability to incorporate sources perturbations that arise in LEO scenarios. More recently, Englander et al [3] and Izzo et al [8] utilized EC to design highly complex mission trajectories. Specifically, Englander et al devised a computational methodology using multi-objective genetic algorithms to perform automated interplanetary mission design, whereas Izzo et al designed a mission trajectory among the Galilean moons of Jupiter that optimized observational conditions at time of spacecraft at the time of flyby.…”

Section: Introductionmentioning

confidence: 99%

Evolutionary Approach to Lambert’s Problem for Non-Keplerian Spacecraft Trajectories

Hinckley

Hitt

2017

Aerospace

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Abstract: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.

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“…Englander [5] posed the problem of choosing a sequence of fly-bys as an integer programming problem that was solved using a GA approach, and nested DE and particle swarm optimization inside a GA to reproduce solutions to the Galileo and Cassini missions [6]. Izzo et al [11] utilized DE to design a highly complex trajectory among the Galilean moons of Jupiter that optimized observational conditions at time of spacecraft at the time of flyby. However, all of these prior efforts assumed idealized physics for the governing equations of motion and, as such, are not sufficient for trajectory planning for LEO.…”

Section: Introductionmentioning

confidence: 99%

Evolved spacecraft trajectories for low earth orbit

Hinckley

Zieba

Hitt

et al. 2014

Proceedings of the 2014 Annual Conference on Genetic and Evolutionary Computation

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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.

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Search for a grand tour of the jupiter galilean moons

Cited by 28 publications

References 13 publications

A Continuation Method for Converting Trajectories from Patched Conics to Full Gravity Models

A Continuation Method for Converting Trajectories from Patched Conics to Full Gravity Models

Evolutionary Approach to Lambert’s Problem for Non-Keplerian Spacecraft Trajectories

Evolved spacecraft trajectories for low earth orbit

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