Search

Near-Earth Objects (NEOs) are asteroids, comets and meteoroids in heliocentric orbits with perihelion below 1.3 au. Similarly to the population of the Main Asteroid Belt, NEOs are primordial bodies and their study can improve our understanding of the origins of the Solar System. With a catalog of over 30~000 known asteroids and approximately 100 listed short-period comets, the NEO population represents an inventory of exploration targets reachable with significantly lower cost than the objects of the Main Asteroid Belt. In addition, the materials present in these bodies could be used to resupply spacecraft en route to other destinations. The trajectories of past missions to NEOs have been designed with the patched-conics technique supplemented by impulsive and/or low-thrust maneuvers and planetary gravity assist. The transfer times range from some months to a few years, and the close-approach speeds relative to the target have been as high as 10 km/s. The design technique described in this work leverages the invariant structures of the circular restricted three-body problem (CR3BP) to connect the vicinity of the Earth with NEOs in low-eccentricity, low-inclination orbits. The fundamental building blocks are periodic orbits around the collinear points L$_1$ and L$_2$ of the Sun-Earth CR3BP. These orbits are used to generate paths that follow the associated hyperbolic invariant manifolds, exit the sphere of influence of the Earth and reach NEOs on nearby orbits. The strategy is simple, can be applied to depart either a libration point orbit or the vicinity of the Earth, and offers attractive performance features.

This contribution focuses on the design of low-energy transfers between planetary moons and presents an efficient technique to compute trajectories characterized by desirable behaviors in the vicinities of the departure and destination bodies. The method utilizes finite-time Lyapunov exponent maps in combination with the Moon-to-Moon Analytical Transfer (MMAT) method previously proposed by the authors. The integration of these two components facilitates the design of direct transfers between moons within the context of the circular restricted three-body problem, and allows the inclusion of a variety of trajectory patterns, such as captures, landings, transits and takeoffs, at the two ends of a transfer. The foundations and properties of the technique are illustrated through an application based on impulsive direct transfers between Ganymede and Europa. However, the methodology can be employed to assist in the design of more complex mission scenarios, such as moon tours.

Given the interest in future space missions devoted to the exploration of key moons in the solar system and that may involve libration point orbits, an efficient design strategy for transfers between moons is introduced that leverages the dynamics in these multi-body systems. The moon-to-moon analytical transfer (MMAT) method is introduced, comprised of a general methodology for transfer design between the vicinities of the moons in any given system within the context of the circular restricted three-body problem, useful regardless of the orbital planes in which the moons reside. A simplified model enables analytical constraints to efficiently determine the feasibility of a transfer between two different moons moving in the vicinity of a common planet. In particular, connections between the periodic orbits of such two different moons are achieved. The strategy is applicable for any type of direct transfers that satisfy the analytical constraints. Case studies are presented for the Jovian and Uranian systems. The transition of the transfers into higher-fidelity ephemeris models confirms the validity of the MMAT method as a fast tool to provide possible transfer options between two consecutive moons.

Mission designers must study many dynamical models to plan a low-cost spacecraft trajectory that satisfies mission constraints. They routinely use Poincar\'e maps to search for a suitable path through the interconnected web of periodic orbits and invariant manifolds found in multi-body gravitational systems. This paper is concerned with the extraction and interactive visual exploration of this structural landscape to assist spacecraft trajectory planning. We propose algorithmic solutions that address the specific challenges posed by the characterization of the topology in astrodynamics problems and allow for an effective visual analysis of the resulting information. This visualization framework is applied to the circular restricted three-body problem (CR3BP), where it reveals novel periodic orbits with their relevant invariant manifolds in a suitable format for interactive transfer selection. Representative design problems illustrate how spacecraft path planners can leverage our topology visualization to fully exploit the natural dynamics pathways for energy-efficient trajectory designs., Comment: 11 pages, 13 figures, to appear at IEEE VIS 2020

The baseline trajectory proposed for the Gateway is a southern Earth-Moon L2 Near Rectilinear Halo Orbit (NRHO). Designed to avoid eclipses, the NRHO exhibits a resonance with the lunar synodic period. The current investigation details the eclipse behavior in the baseline NRHO. Then, phase control is added to the orbit maintenance algorithm to regulate perilune passage time and maintain the eclipse-free characteristics of the Gateway reference orbit. A targeting strategy is designed to periodically target back to the long-horizon virtual reference if the orbit diverges over time in the presence of additional perturbations.

The upcoming Lunar IceCube (LIC) mission will deliver a 6U CubeSat to a low lunar orbit via a ride-share opportunity during NASAs Artemis 1 mission. This presents a challenging trajectory design scenario, as the vast change in energy required to transfer from the initial deployment state to the destination orbit is compounded by the limitations of the LICs low-thrust engine. This investigation addresses these challenges by developing a trajectory design framework that utilizes dynamical structures available in the Bicircular Restricted Four-Body Problem (BCR4BP) along with a robust direct collocation algorithm. Maps are created that expedite the selection of invariant manifold paths from a periodic staging orbit in the BCR4BP that offer favorable connections between the LIC transfer phases. Initial guesses assembled from these maps are passed to a direct collocation algorithm that corrects them in the BCR4BP while including the variable low-thrust acceleration of the spacecraft engine. Results indicate that the ordered motion provided by the BCR4BP and the robustness of direct collocation combine to offer an efficient and adaptable framework for designing a baseline trajectory for the LIC mission.

Spacecraft departing from the Gateway in a Near Rectilinear Halo Orbit (NRHO) experience gravitational forces from the Moon, the Earth, and the Sun, all of which can be simultaneously significant. These complex dynamics influence the post-separation risk of recontact with the Gateway and the eventual destinations of the departing spacecraft. The current investigation examines the flow of objects leaving NRHOs in the Bicircular Restricted Four-Body Problem, and results are applied to heliocentric escape and lunar impact trajectories in a higher-fidelity ephemeris model. Separation maneuver magnitude, direction, and location are correlated with risk of recontact with the Gateway and successful departure to various destinations.

A proposed Gateway facility in a lunar Near Rectilinear Halo Orbit (NRHO) will serve as an outpost in deep space, with spacecraft periodically arriving and departing. Departing objects will include logistics modules, requiring safe disposal, cubesats, deployed to various destinations, and debris objects, whose precise paths may be unknown. Escape dynamics from NRHOs are complex; primarily influenced by the Earth and Moon within the orbit, spacecraft are significantly impacted by solar gravity upon departure. The current investigation explores the dynamics of departure from the NRHO, including the risk of debris recontact, safe heliocentric disposal, and deployment to select destinations.