Search

Autonomy is an increasingly essential component of future space missions, and new technologies are necessary to accommodate off-nominal occurrences onboard that may pose risks to mission success. Identifying a suitable maneuver plan for onboard, low-thrust mission applications in cislunar space remains challenging, particularly in the case of unanticipated events. This investigation addresses this challenge by demonstrating artificial neural networks as promising tools for accurately estimating startup solutions for a conventional targeting-based iterative guidance and control algorithm. This blended approach produces a robust 'hybrid' architecture that simultaneously benefits from the computational simplicity of neural networks and the robustness of differential corrections to satisfy mission requirements. In this paradigm, a multiple shooting scheme is incorporated directly into a reinforcement learning process, tasking the resulting neural network controller with convergence basin identification for trajectory recovery. Rapid low-thrust maneuver planning is demonstrated in a 'runaway' spacecraft scenario, where deviation over time from a planned near rectilinear halo orbit path renders stationkeeping ineffective and demands an alternative approach to determine an effective recovery plan. The proposed Neural Network-Initialized Targeting (NNIT) algorithm significantly extends the recovery window compared to a crossing-control orbit maintenance approach and exhibits promising results in identifying both the timing and control components for low-thrust maneuver plans. • Neural networks are leveraged to quickly identify convergence basins for low-thrust. • Reinforcement learning enables a cislunar autonomous trajectory recovery scenario. • Onboard neural network risk is mitigated by introducing a hybrid G&C architecture. • Timing and control components of low-thrust maneuver plans are considered. • Recoverable window is extended by three weeks compared to a stationkeeping method. [ABSTRACT FROM AUTHOR]

A proposed Gateway facility in a lunar near rectilinear halo orbit (NRHO) will serve as an outpost in cislunar space, with various spacecraft periodically arriving and departing. The dynamics in the proximity of an NRHO are complex; the motion is primarily influenced by the Earth and Moon within the vicinity of the NRHO, but spacecraft are significantly impacted by solar gravity upon departure from the lunar vicinity. The current investigation explores two aspects of the trajectory dynamics after a separation maneuver from Gateway: departure dynamics and escape dynamics. The departure dynamics analysis focuses on the immediate fate of the trajectory after a maneuver, while the object remains close to the NRHO. The escape dynamics analysis details the outcome of the trajectory, that is, whether it remains in the Earth-Moon vicinity or escapes to heliocentric space. This investigation identifies a range of parameters and metrics at the different stages along the departed trajectory that inform the trajectory design process. [ABSTRACT FROM AUTHOR]

Near Rectilinear Halo Orbits (NRHOs) are stable or nearly stable orbits that are defined as part of the L1 and L2 halo orbit families in the circular restricted three-body problem. Within the Earth-Moon regime, the L1 and L2 NRHOs are proposed as long horizon trajectories for cislunar exploration missions, including NASA's upcoming Gateway mission. The spacecraft, however, incurs continuous deviations due to unmodeled forces and orbit determination errors in this dynamically sensitive region. The current investigation focuses on an impulsive stationkeeping technique to maintain the spacecraft near a virtual reference orbit despite these uncertainties. The flow dynamics in the region are utilized to identify appropriate maneuver and target locations. The investigation reflects the impact of various factors on maneuver cost and efficacy. For orbits where position and velocity states are particularly sensitive to epoch time, an additional feedback control strategy is applied for phasing constraints. [ABSTRACT FROM AUTHOR]

Path planning in the circular restricted 3-body problem (CR3BP) is frequently guided by the forbidden regions and manifold arcs. However, when low thrust is employed to modify the spacecraft trajectory, these dynamical structures pulsate with the varying Hamiltonian value. In a combined CR3BP, low-thrust (CR3BP + LT) model, an additional low-thrust Hamiltonian is available that remains constant along low-thrust arcs given suitable assumptions. Accordingly, the analogous low-thrust forbidden regions and manifolds remain static and are useful guides for low-thrust trajectory design. Strategies leveraging these structures and other insights from the CR3BP + LT are explored to construct transit and capture itineraries. [ABSTRACT FROM AUTHOR]

Mission designers must study many dynamical models to plan a low-cost spacecraft trajectory that satisfies mission constraints. They routinely use Poincare 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. [ABSTRACT FROM AUTHOR]

In the near future, several space applications in the Earth-Moon system may require a spacecraft to hold a stable motion, but the transfer trajectory infrastructure to access such stable motions has not been fully investigated yet. The triangular libration points, L 4 and L 5 , in the Earth-Moon system have long been thought of as potential locations for a communications satellite. Recently, Distant Retrograde Orbits (DROs) and Near-Rectilinear Halo Orbits (NRHOs) in the vicinity of the Moon have been identified as orbits of interest for manned and unmanned missions with a focus on operations in cislunar space. The triangular libration points, as well as lunar DROs and NRHOs describe special types of possible motion for a spacecraft/satellite that is influenced solely by the gravitational fields of the Earth and the Moon. What is common to the three types of solutions is that they are practically stable, that is, a spacecraft/satellite can naturally follow the solution for extended periods of time without requiring significant course adjustment maneuvers. This investigation proposes the lunar region as the central link to a transfer network that enables travel throughout the Earth-Moon system, connecting the lunar region to the vicinity of the Earth and the neighborhood of the triangular libration points. The work presented here also contributes to the infrastructure supporting such a network by expanding the transfer options available between these regions. Several new transfer options between regions of stability are presented and discussed, including transfer options between Low Earth Orbit (LEO) and lunar DRO, lunar DRO and periodic orbits near L 4 and L 5 , as well as lunar DRO and L 2 NRHOs. Underlying dynamical mechanisms enabling transfers between selected orbits are analyzed, and sample itineraries are provided. [ABSTRACT FROM AUTHOR]

The motion of a spacecraft in proximity to a small body is significantly perturbed due to its irregular gravity field and solar radiation pressure. In such a strongly perturbed environment, the coupling effect of the orbital and attitude motions exerts a large influence that cannot be neglected. However, natural orbit-attitude coupled dynamics around small bodies that are stationary in both orbital and attitude motions have yet to be observed. The present study therefore investigates natural coupled motion that involves both a Sun-synchronous orbit and Sun-tracking attitude motion. This orbit-attitude coupled motion enables a spacecraft to maintain its orbital geometry and attitude state with respect to the Sun without requiring active control. Therefore, the proposed method can reduce the use of an orbit and attitude control system. This paper first presents analytical conditions to achieve Sun-synchronous orbits and Sun-tracking attitude motion. These analytical solutions are then numerically propagated based on non-linear coupled orbit-attitude equations of motion. Consequently, the possibility of implementing Sun-synchronous orbits with Sun-tracking attitude motion is demonstrated. [ABSTRACT FROM AUTHOR]

Trajectory design increasingly leverages multi-body dynamical structures that are based on an understanding of various types of orbits in the Circular Restricted Three-Body Problem (CR3BP). Given the more complex dynamical environment, mission applications may also benefit from deeper insight into the attitude motion. In this investigation, the attitude dynamics are coupled with the trajectories in the CR3BP. In a highly sensitive dynamical model, such as the orbit-attitude CR3BP, periodic solutions allow delineation of the fundamental dynamical structures. Periodic solutions are also a subset of motions that are bounded over an infinite time-span (assuming no perturbing factors), without the necessity to integrate over an infinite time interval. Euler equations of motion and quaternion kinematics describe the rotational behavior of the spacecraft, whereas the translation of the center of mass is modeled in the CR3BP equations. A multiple shooting and continuation procedure is employed to target orbit-attitude periodic solutions in this model. Application of Floquet theory and Poincaré mappings to identify initial guesses for the targeting algorithm are described. In the Earth–Moon system, representative scenarios are explored for axisymmetric vehicles with various inertia characteristics, assuming that the vehicles move along L 1 / L 2 Lyapunov orbits as well as distant retrograde orbits. A rich structure of possible periodic behaviors appears to pervade the solution space in the coupled problem. The stability analysis of the attitude dynamics for the available families is included. Among the computed solutions, marginally stable and slowly diverging rotational behaviors exist and may offer interesting mission applications. [ABSTRACT FROM AUTHOR]

To investigate the behavior of a spacecraft near a pair of irregular bodies, consider a three body configuration (one massless). Two massive bodies, P 1 and P 2 , form the primary system; each primary is initially modeled as an ellipsoid. Concepts and tools similar to those applied in the Circular Restricted Three-Body Problem (CRTBP) are exploited to construct bounded trajectories for a third body in such a synchronous system. Then, to accommodate model uncertainties and other non-gravitational perturbations, a strategy to maintain a spacecraft near a reference path is proposed. [ABSTRACT FROM AUTHOR]

The attitude motion of a rigid spacecraft in fully nonlinear reference orbits is explored in this investigation, within the context of the planar circular restricted three-body problem. The reference trajectories originate from the Lyapunov families about the Lagrangian points L, and L2 in the Earth-Moon system. Kane's method is employed to derive the equations of motion, which are numerically solved. The capability to reproduce the dynamic response is leveraged to understand the attitude behavior across the Lyapunov families. The problem formulation and the simulation environment are detailed. The inertia properties of the spacecraft are varied across the periodic family of orbits. Finally, attitude maps are introduced to summarize the results and identify the regions, in terms of the orbit size and inertia properties, where the spacecraft maintains the initial orientation with respect to the rotating frame in the circular restricted three-body problem. [ABSTRACT FROM AUTHOR]

This investigation explores the natural dynamics in a multi-body regime for formation flying applications. Natural regions that are suitable to maintain multiple spacecraft in a loose formation are determined in an inertial coordinate frame. Locating a formation of spacecraft in these zones leads to a smaller variation in the mutual distance between the spacecraft and the pointing direction of the formation. These suitable regions approximate quadric surfaces along a reference trajectory and the relationship between the dynamical evolution of the quadric surfaces and the eigenstructure of the reference trajectory is examined. [ABSTRACT FROM AUTHOR]

Recent investigations of trajectory options that incorporate solar sails have been motivated by missions to observe planetary poles or to communicate with an outpost at the lunar south pole. Designing reference trajectories and understanding their fundamental dynamics are the necessary first steps toward flying spacecraft in dynamically complicated regimes. However, the existence of a reference orbit alone is insufficient for flight operations. Two variations of a turn-and-hold strategy are examined for flight-path control: an approach that implements multiple turns to achieve a target in an error-free scenario and an approach that incorporates a look-ahead strategy to accommodate representative errors. [ABSTRACT FROM PUBLISHER]

Abstract: Recent discoveries by the Cassini spacecraft have generated interest in future missions to further explore the moons of Saturn as well as other small bodies in the solar system. Incorporating multi-body dynamics into the preliminary design can aid the design process and potentially reduce the cost of maneuvers that are required to achieve certain objectives. The focus in this investigation is the development and application of additional design tools to facilitate preliminary trajectory design in a multi-body environment where the gravitational influence of both primaries is quite significant. Within the context of the circular restricted 3-body problem, then, the evolution of trajectories in the vicinity of the smaller primary (P2) that are strongly influenced by the distant larger primary (P1) is investigated. By parameterizing the orbits in terms of radius and periapse orientation relative to the P1–P2 line, the short- and long-term behaviors of the trajectories are predictable. Initial conditions that yield a trajectory with a particular set of desired characteristics are easily selected from periapsis Poincaré maps for both short- and long-term orbits. Analysis in the Sun–Saturn and Saturn–Titan systems serves as the basis for examples of mission design applications. [Copyright &y& Elsevier]

Abstract: Upcoming National Aeronautics and Space Administration (NASA) mission concepts include satellite arrays to facilitate imaging and identification of distant planets. These mission scenarios are diverse, including designs such as the terrestrial planet finder-occulter (TPF-O), where a monolithic telescope is aided by a single occulter spacecraft, and the micro-arcsecond X-ray imaging mission (MAXIM), where as many as 16 spacecraft move together to form a space interferometer. Each design, however, requires precise reconfiguration and star tracking in potentially complex dynamic regimes. This paper explores control methods for satellite imaging array reconfiguration in multi-body systems. Specifically, optimal nonlinear control and geometric control methods are derived and compared to the more traditional linear quadratic regulators, as well as input state feedback linearization. These control strategies are implemented and evaluated for the TPF-O mission concept. [Copyright &y& Elsevier]

Libration point orbits may be advantageous locations for spacecraft imaging formations. Therefore, control of these arrays in multi-body regimes is critical. Ideally, the motion of spacecraft within an imaging array minimizes fuel usage while maximizing resolution of a distant object in some mission-specified frequency band. Maximizing image resolution implies adequate coverage of the (u, v) plane in a specified observation time. In the current work, an optimization problem to minimize fuel usage while maximizing an image metric is formulated in the circular restricted three-body problem. Optimization is simplified by subdividing the task into two distinct smaller problems: imaging optimization and fuel optimization. Then, these problems are solved separately. Image reconstruction and coverage of the (u. v) plane are simulated for interferometric spacecraft configurations, demonstrating potential applications of the algorithm and the resulting motion. [ABSTRACT FROM AUTHOR]

The general objective is the development of efficient techniques for preliminary design of trajectory arcs in nonlinear autonomous dynamic systems in which the desired solution is subject to algebraic interior and/or exterior constraints. For application to the n-body problem, trajectories must satisfy specific requirements, e.g., periodicity in terms of the states, interior or boundary constraints, and specified coverage. Thus, a strategy is formulated in a sequence of increasingly complex steps: 1) a trajectory is first modeled as a series of arcs (analytical or numerical) and general trajectory characteristics and timing requirements are established; 2) the specific constraints and associated partials are formulated; 3) a corrections process ensures position and velocity continuity while satisfying the constraints; and finally, 4) the solution is transitioned to a full model employing ephemerides. Though the examples pertain to spacecraft mission design, the methodology is generally applicable to autonomous systems subject to algebraic constraints. For spacecraft mission design applications, an immediate advantage of this approach, particularly for the identification of periodic orbits, is that the startup solution need not exhibit any symmetry to achieve the objectives. [ABSTRACT FROM AUTHOR]

Abstract: One immediate and cost-effective solution for lunar south pole coverage is a low-thrust pole-sitter spacecraft. When the spacecraft''s fuel is nearly expended, it can also be used as part of a larger constellation that might require more time to deploy. This study examines the feasibility of the lunar pole-sitter from a dynamical standpoint. The model includes the effects of Earth oblateness, solar wind, gravity perturbations, and shadowing on a 500kg spacecraft equipped with an NSTAR thruster. Lunar librations are also incorporated into path constraints on elevation angle and altitude from a potential ground station at the Shackleton crater. The solutions utilize multiple phases, including transfer-out from the International Space Station and spiral-down into a stable lunar orbit once the pole-sitting period is completed. Computation of the trajectory is based on a 12th-order collocation scheme. A direct transcription routine maximizes the time in the coverage phase, increasing the time span from 387 to 451 days. During this period, the minimum elevation angle relative to the Shackleton crater site is 6°. [Copyright &y& Elsevier]

Potential end-of-mission scenarios to be considered for the Cassini spacecraft must satisfy planetary quarantine requirements designed to prevent contamination of a pristine environment, which could include Titan and the other Saturnian moons. One assumed acceptable option for safe disposal of the spacecraft includes Saturn impact trajectories. Two classes of impact trajectories are investigated: short-period orbits characterized by periods of 6- 10 days and long-period orbits with periods greater than 850 days. To impact Saturn with short-period orbits, a series of successive Titan flybys is required to increase inclination and decrease periapsis to within Saturn's atmosphere, while simultaneously avoiding the rings and mitigating A V expenditures. To ensure that the spacecraft is not prematurely damaged by material in the rings, Tisserand graphs are employed to determine when the ring-plane crossing distance is within the F-G ring gap: the necessary geometry for the penultimate transfer. For long-period impact trajectories, solar gravity is exploited to significantly lower periapsis. Depending on the size and orientation of the long-period orbit, a maneuver (<50 m/s) at apoapsis must be added to ensure impact. For sufficiently large orbits with favorable characteristics, solar gravity alone drops the spacecraft's periapsis into Saturn's atmosphere. No maneuver is necessary after the final Titan flyby, providing an attractive "flyby-and-forget" option. [ABSTRACT FROM AUTHOR]