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For high-resolution imaging implementations, the spacecraft attitude tracking control accuracy is crucial to determining the imaging quality. This investigation addresses the attitude tracking issue of imaging spacecraft subject to system uncertainties (unavailable inertia tensor, unexpected disturbances, and actuator faults). An adaptive sliding mode control (SMC) strategy is proposed to guarantee practical fixed-time closed-loop stability even in the presence of system uncertainties. Unlike existing methodologies, the sliding mode surface is developed to satisfy a novel sufficient condition of fixed-time stability. The sliding manifold design also circumvents the unwinding phenomenon arising in quaternion representations. Particularly, this controller is developed to generate a smooth control profile by using a new parameter update law. Rigorous Lyapunov analyses are further employed to ensure the fixed-time closed-loop stability irrespective of the system initial states. Finally, numerical examples are performed to demonstrate the feasibility and highlight the inherent features of the derived control law.

This paper presents an indirect estimation based active disturbance rejection control (ADRC) for solar sail station-keeping. In contrast to previous applications of ADRC to spacecraft station-keeping the proposed approach estimates the deviation of the orbit state from the reference orbit, instead of estimating the orbit state directly. Since the deviation of the orbit state varies more slowly than the state itself, the indirect estimation significantly improves the estimation accuracy and hence the station-keeping performance. Moreover, this approach enables a large control time step which is two orders of magnitude larger than those in previous applications. This is beneficial to the implementation of attitude control and power consumption in practice. In addition, errors in the optical properties of the solar sail are taken into account which is a major challenge for solar sail station-keeping that has not been considered in previous station-keeping control design. The simulation results show that the presented ADRC station-keeping is robust to errors in these optical properties.

High tumbling rates of uncooperative target pose strong technical challenges and collision risks that can prevent removal of the debris using contact, such as with robotic arms or capture nets. Ele...

Current control approaches for solar sail station-keeping on libration point orbits have not considered the degradation of the sail’s optical properties. However, significant optical degradation could lead to poor station-keeping performance or even complete failure. This paper presents an integrated guidance and control strategy to address this problem by updating the reference orbit based on in situ estimation. An exponential optical degradation model is incorporated into the solar radiation acceleration model, and an on-line reference orbit update approach is incorporated into the station-keeping, coupled with an active disturbance rejection controller. The reflection coefficient is estimated on-line and the reference orbit is updated discretely when the optical properties have degraded by a prescribed amount. This strategy provides discrete updates to the reference orbits such that the perturbation due to the optical degradation is maintained within a small range. These smaller perturbations can be dealt with by the controller’s robustness and station-keeping can be sustained for long durations even in the presence of large optical degradation.

This paper addresses the attitude control of nano-spacecraft with three reaction wheels and three magnetorquers in low Earth orbits, and develops algorithms to maintain control when any one of the reaction wheels completely fails. In the event of a reaction wheel failure, it is shown that the key challenge of maintaining accurate autonomous three-axis tracking is a problem of singularity avoidance in control allocation. In particular, control law singularities occur depending on the type of orbit, the environment, and the attitude, where the system is no longer controllable. Two algorithmic singularity avoidance techniques based on singularity prediction and artificial potential functions are proposed. In addition, an optimized reaction wheel configuration is proposed to enhance singularity avoidance capability. To illustrate the effectiveness of the singularity avoidance control laws, they are tested in simulations that include the effects of perturbing environmental torques, reaction wheel jitter, and actuator misalignment.

On-orbit assembly strategies are proposed for an orb-shaped solar array which forms the main structure of a space solar power station. A dynamic model for the structural vibration under the influence of the gravity gradient is proposed by combining the Tschauner-Hempel equation and a finite element model of an elastic beam. The motion of the robot, used for on-orbit assembly, is described analogously to a classical vehicle-bridge coupled dynamic model. Assembly strategies are presented and analyzed, consisting of a single-robot assembly strategy, an assembly strategy using two robots and structural reinforcements. The proposed methods are verified in simulation and tuned to minimize the structural vibration during assembly.

A supercavitating underwater vehicle uses gas or vapor to reduce its skin friction enabling it to travel up to speeds of 200 m/s. However, traveling at extreme speeds can be challenging for control design since sudden disturbances or faults, even for a short time, can lead to large deviations from the reference trajectory. This paper addresses the problem of controlling a supercavitating vehicle which incorporates the actuator constraints and loss of effectiveness of the cavitator. An adaptive control law is designed based on the backstepping control method which includes an adaptive fault-tolerant component. The bounded adaptive update law can estimate the upper bound of the unknown total uncertainties including those induced by actuator faults. The proposed control is shown to guarantee stability in the presence of faults by using Lyapunov theory.

This paper addresses the problem of optimally repointing the optical axis of a spacecraft to align with the target direction. A new metric defining the repointing error is proposed where the corresponding kinematic equations provide a simple and convenient form for control design. The proposed control integrates a Control Lyapunov Function (CLF) approach with a sliding mode controller which simultaneously guarantees the optimality and robustness of the closed-loop system. Firstly, a CLF based control scheme is used to ensure that the state optimally converges to the sliding surface. Then a fixed-time non-singular terminal sliding mode controller is employed to provide robust convergence to the origin along the sliding surface. The convergence time is finite for any initial states and is thus useful for applications with critical time constraints. The region of attraction and convergence time is analyzed. Finally, numerical investigations are conducted to verify the effectiveness and superiority of the proposed algorithm with respect to the classical CLF method.

This paper describes the trajectory tracking of an underactuated autonomous underwater vehicle (AUV) with three control inputs (surge, yaw and pitch moment) that operates in the presence of time-varying disturbances. The AUV kinematics are described in global coordinates as a Hamiltonian system on the Lie group SE(3) and the boundary-value problem arising from the geometric framing of Pontryagin’s Maximum Principle is applied to the vehicle kinematics. This 6-dimensional boundary value problem is solved using a numerical shooting method and a novel semi-analytical Lie group integrator. The integrator uses Rodrigue’s formula to express the exact solution of the solution curves, is symplectic and preserves energy and momentum. Inverse dynamics is applied to construct an inner-loop controller, which accounts for constraints on maximum torque and force via time reparametrization. This inner-loop control, which would drive the AUV along the reference trajectory in perfect conditions, is combined with a disturbance observer to construct an outer-loop controller, which ensures stability in the presence of bounded disturbances. Simulation results complete the work.

In geostationary orbit, long eclipses and the seasonal variations in the direction and intensity of the solar input can cause damage to sensitive equipment during attitude maneuvers, which may inadvertently point the equipment towards the Sun. The requirement that transmitting and receiving antennae remain pointed towards the Earth creates further restrictions to pointing directions. The aim of the study is to construct a novel geometric and reinforcement-learning-based method to determine attitude guidance maneuvers that maintain the equipment in safe and operational orientations throughout an attitude maneuver. The attitude trajectory is computed numerically using the geometric framing of Pontryagin’s maximum principle applied to the vehicle kinematics using the global matrix Lie group representation on SO(3), and the angular velocities are shaped using free parameters. The values of these free parameters are determined by a reinforcement learning algorithm to avoid the forbidden areas while maintaining the pointing in operational areas (modeled as subsets of the two-sphere of all possible pointing directions of a particular axis). The method is applied to a model geosynchronous satellite and demonstrated in a simulation.

A low-jerk attitude guidance method is developed, based on an analytical smoothing of a bang-off-bang maneuver. A set of closed-form equations are derived and used to plan constrained low-jerk maneuvers, with prescribed boundary conditions, inertia, time and maximum control torque. The guidance law is first developed for one-dimensional and then three-dimensional rotations using two different approaches: (i) by designing a rotation about the Euler axis and (ii) by using the inverse kinematics equations. A generic model of the torque induced by a multi-body appendage is derived using the lumped-parameter method. This method can also be used to approximate the dynamic behavior of flexible appendages. The simulations results show that the smoothing techniques reduce the excitation of multi-body and flexible structures during a slew maneuver.

This paper addresses the optimal orbit rendezvous problem for thrust-vectoring spacecraft. In this scenario, the chaser spacecraft is equipped with only one impulsive thruster fixed in the body frame, while the thrust direction is pointed using attitude control of the spacecraft. An improved particle swarm optimization (PSO) algorithm is employed to generate an optimal multiple-burn rendezvous trajectory in terms of thruster pulse duration and required attitude. A coordinate-free, finite-time, attitude control is developed to ensure that this required thrust vector is met exactly at the prescribed time. The proposed orbit-attitude guidance algorithm is compared with one which utilizes a conventional PSO and is shown to significantly improve performance. In addition, the set-point attitude control is illustrated, in simulation, to deliver the required thrust direction at the prescribed time.

Solar sail halo orbits designed in the Sun-Earth circular restricted three-body problem (CR3BP) provide inefficient reference orbits for station-keeping since the disturbance due to the eccentricity of the Earth’s orbit has to be compensated for. This paper presents a strategy to compute families of halo orbits around the collinear artificial equilibrium points in the Sun-Earth elliptic restricted three-body problem (ER3BP) for a solar sail with reflectivity control devices (RCDs). In this non-autonomous model, periodic halo orbits only exist when their periods are equal to integer multiples of one year. Here multi-revolution halo orbits with periods equal to integer multiples of one year are constructed in the CR3BP and then used as seeds to numerically continue the halo orbits in the ER3BP. The linear stability of the orbits is analyzed which shows that the in-plane motion is unstable while the out-of-plane motion is neutrally stable and a bifurcation is identified. Finally, station-keeping is performed which shows that a reference orbit designed in the ER3BP is significantly more efficient than that designed in the CR3BP, while the addition of RCDs improve station-keeping performance and robustness to uncertainty in the sail lightness number.

Electromagnetic docking could enable autonomous spacecraft docking with no need for propellant consumption and without plume contamination. This paper addresses the robust electromagnetic docking problem for spacecraft in the presence of external disturbances, fault signals, unknown mass, elliptical eccentricity, measurement errors and input constraints. In this scenario, an intermediate observer is developed to estimate the relative motion information and the lumped disturbance resulting from these uncertainties. Based on this, an anti-disturbance controller is proposed, where the compensation of the lumped disturbance is considered. It is proved via Lyapunov analysis that the intermediate observer-based controller can achieve the objective of spacecraft electromagnetic docking with input constraints and in the presence of uncertainties. Finally, the observer-based controller is illustrated, in simulation, to demonstrate the effectiveness and improved performance compared with a disturbance observer-based controller.

This article addresses the coupled orbit-attitude tracking problem using low-thrust propulsion while aiming to minimize on-board spacecraft system communication. An adaptive controller is proposed by employing an event-triggered control and an extended-state-observer, where a simple strategy to tune the observer parameters is provided. Moreover, the event-triggering strategy updates and allocates the control signal to the thrusters at prescribed discrete events and is shown to significantly reduce the data-rate requirement. Finally, the performance of the controller is illustrated through numerical examples.

The capture or docking of a spacecraft with a target spacecraft is an important component of future space debris capture and orbit servicing missions. One proposed concept includes using a robotic arm mounted on the chaser spacecraft to connect and manipulate the target's motion. This paper focuses on the post-capture attitude control of the target which will bring the target to rest and then control it in a prescribed way. The attitude control can be achieved by designing a robust control in the known chaser reference frame and treating the dynamics due to the uncertain inertia as an unknown disturbance. However, this can lead to poor performance since this uncertain torque can be large. In this paper, it is shown that the in-situ construction of the chaser-target dynamics can lead to improved performance. Based on the combined re-constructed model, a new disturbance observer based control with robust dynamic control allocation is developed. This approach is robust to the uncertainties in the re-constructed model and enables the prescription of convergence and stability properties. The proposed approach is demonstrated through numerical simulation of an actively controlled 12U CubeSat that has captured an inactive 12U CubeSat spacecraft.

This paper tackles the problem of optimal attitude control using a minimal number of attitude thrusters. Three possible control solutions to this problem are presented: 1) a logic-based controller ...

A minimal actuator design is proposed for orbit-attitude control of a spacecraft. The actuator design includes a single thruster, fixed in the body frame, and a variable-speed control moment gyrosc...

A global motion planning method is described based on the solution of minimum energy-type curves on the frame bundle of connected surfaces of arbitrary constant cross sectional curvature ϵ . Applying the geometric framing of Pontryagin’s principle gives rise to necessary conditions for optimality in the form of a boundary value problem. This arbitrary dimensional boundary value problem is solved using a numerical shooting method derived from a general Lax pair solution. The paper then specializes to the 3-dimensional case where the Lax pair equations are integrable. A semi-analytic method for matching the boundary conditions is proposed by using the analytic form of the extremal solutions and a closed form solution for the exponential map. This semi-analytical approach has the advantage that an analytic description of the control accelerations can be derived and enables actuator constraints to be incorporated via time reparametrization. The method is applied to two examples in space mechanics: the attitude control of a spacecraft with two reaction wheels and the spacecraft docking problem.

This paper extends a linear quadratic regulator-based control scheme for station-keeping on libration point orbits to enable continued tracking in the event of thruster failure. In the first instan...

This paper develops a quaternion attitude tracking control with an adaptive gains parameter that can be tuned to compensate for disturbances with known bound. The adaptive gain is described by a simple, but singular, differential equation and the corresponding adaptive control is shown to asymptotically track a reference attitude. However, this control requires the bound on the disturbance torque to be known in order to appropriately tune the controller to compensate for it. Using a linear state observer to estimate the disturbance torque and compensating for the disturbance at each sampling period the adaptive control can achieve asymptotic tracking in the presence of an unknown disturbance torque. In this case the error in the estimation, rather than the entire disturbance, is compensated for by the adaptive gain at each sampling period. Simulations demonstrate that an improved tracking performance can be achieved when compared to standard quaternion tracking controls.

This paper addresses the guidance and tracking problem for a rigid-spacecraft using two reaction wheels (RWs). The guidance problem is formulated as an optimal control problem on the special orthog...

Repointing maneuvers of a spacecraft in staring mode are investigated where the optical axis is required to align with the target orientation. Different from traditional three-axis reorientation maneuvers, the rotation about the optical axis is free of constraints for repointing maneuvers. Both static target observation and moving target detection constraints are considered. The problem is then formulated as a finite-time horizon optimal control problem with nonlinear terminal constraints. A simple and efficient state-dependent Riccati equation(SDRE) based dynamic programming approach is applied to tackle this nonlinear optimal control problem. The convergence of the attitude from initial conditions to the desired terminal constraint is rigorously proved for the first time. Considering the inability of the SDRE method to deal with the problem of large angle maneuvers, an improved SDRE approach combined with a waypoint is proposed to enhance control performance. Finally, numerical investigations are conducted and compared with the real optimal solutions obtained by using the optimization software.

This paper addresses the relative position tracking and attitude synchronization control problem for spacecraft formation flying (SFF). Based on the derived relative coupled six-degree-of-freedom dynamics, a robust adaptive finite-time fast terminal sliding mode controller is proposed to achieve the desired formation in the presence of model uncertainties and external disturbances. It is shown that the designed controller is effective for changing information exchange topology making it robust to node failure. Then, the artificial potential function method is employed to generate collision avoidance schemes to modify the controller such that inter-agent collision avoidance can be ensured during the formation maneuver, which is critical for practical missions. The stability of the overall closed-loop system is proved by using Lyapunov theory. Finally, numerical examples for a given SFF scenario are presented to illustrate the performance of the controller.

The Lunar Meteoroid Impact Observer (LUMIO) is one of the four projects selected within ESA’s SysNova competition to develop a small satellite or scientific and technology demonstration purposes to be deployed by a mothership around the Moon. Themission utilizes a 12U form-factor CubeSat which carries the LUMIO-Cam, an optical instrument capable of detecting light flashes in the visible spectrum to continuously monitor and process the meteoroids impacts. In this chapter, we will describe the mission concept and focus on the performance of a novel navigation concept using Moon images taken as byproduct of the LUMIOCam operations. This new approach will considerably limit the operations burden on ground, aiming at autonomous orbit-attitude navigation and control. Furthermore, an efficient and autonomous strategy for collection, processing, categorization, and storage of payload data is also described to cope with the limited contact time and downlink bandwidth. Since all communications have to go via a lunar orbiter, all commands and telemetry/data will have to be forwarded to/from the mothership. This will prevent quasi-real-time operations and will be the first time for CubeSats as they have never flown without a direct link to Earth. This chapter was derived from a paper the authors delivered at the SpaceOps 2018 conference.

This paper presents an approach to motion planning for left-invariant kinematic systems defined on the six-dimensional frame bundles of symmetric spaces of constant cross-sectional curvature. A covering map is used to convert the original differential equation into two coupled equations each evolving on a three-dimensional Lie group. These lower dimensional systems lend themselves to a minimal global representation that avoids singularities associated with the use of exponential coordinates. Open-loop and closed-loop kinematic control problems are addressed to demonstrate the use of this mapping for analytical and numerical motion planning methods. The approach is applied to a spacecraft docking problem using two different types of actuation: 1) a fully actuated continuous low-thrust propulsion system; and 2) an underactuated single impulsive thruster and reaction wheel system.

Robust, co-ordinate free, thrust-vector tracking controls are designed for a single Variable Speed Control Moment Gyro (VSCMG). The proposed controls, based on an adaptive active disturbance rejection control and higher-order sliding mode control are presented and compared. These controls guarantee the closed-loop pointing error converges to and remain within a bounded region of the zero error even in the presence of the variable inertia induced by the VSCMG, thruster off-set and uncertain disturbance torques. Secondly, the pointing control law is developed to address the problem of spacecraft rendezvous using only a single thruster and a single VSCMG. Here a pointing guidance law is developed that weights the desired thrust of the rendezvous using a scaling function of the pointing error. This guidance law is shown to significantly improve the performance of the rendezvous maneuver with respect to a guidance law that does not utilize the scaling function. Simulations of a rendezvous maneuver with a 12 U CubeSat are undertaken to demonstrate the applicability of the developed control laws.

This paper addresses the problem of optimally repointing the optical axis of a spacecraft to align with a target. A new metric defining the repointing error is proposed where the corresponding kinematic equations provide a simple and convenient form for control design. A control is proposed that integrates a Control Lyapunov Function (CLF) approach with a sliding mode controller which guarantees both optimality and robustness of the closed-loop system. Moreover, a CLF based control scheme ensures that the state optimally converges to a sliding surface, while a fixed-time non-singular terminal sliding mode controller provides robust convergence to the origin while on this surface. The region of attraction and convergence time is analyzed. Finally, numerical investigations are conducted to verify the effectiveness and superiority of the proposed algorithm with respect to the classical CLF method and an integral sliding mode based CLF approach.

Recently there has been a resurgence of interest in missions to the moon and a major challenge of such missions is to provide a continuous communication between the Earth and the Moon's far side. Orbits around the $L_2$ Earth-Moon Lagrange point have been a topic of interest in this field due to their potential for constant communication with both the Earth and the Moon, however the Lagrange point orbits are innately unstable and so station-keeping control is required to maintain them. Station-keeping problems are highly nonlinear and a traditional approach to control design is first to linearize the nonlinear system. However, this first-order approximation introduces errors if there are large injection errors. This paper demonstrates how a simple Extended State Observer (ESO) can be used to improve the convergence time of spacecraft to the reference orbit given with large injection errors. Additionally, solar radiation pressure (SRP) a dominant disturbance in deep-space, can lead to inefficient station-keeping if it is not taken into account in the reference orbit design. New reference orbits can be designed that exploit the SRP perturbation but this assumes that it is known apriori. Here we show how an ESO could provide an in-orbit measurement of the SRP which could be used to modify the reference trajectory to a more fuel efficient one. Finally, it is shown how an ESO can be used to estimate, not only the disturbance, but simultaneously the velocity of the spacecraft meaning that only the position of the spacecraft is required.

This paper describes a global motion planning method for vehicles with actuator constraints based on the semi-analytic solution of minimum energy-type curves on the frame bundle of connected surfaces of arbitrary cross sectional curvature. The method is semi-analytic and solves the boundary-value problem arising from the geometric framing of the Pontryagin's maximum principle applied to the vehicle kinematics where the velocities are defined analytically in terms of three parameters. Numerical shooting on an iterative Lie group expression of the curve in the group is then employed to match the group boundary conditions. This approach has the advantage that an analytic description of the control accelerations can be derived and enables actuator constraints to be incorporated via time reparametrization. The method is applied to a practical example from space mechanics, the spacecraft docking problem with actuator constraints.

This paper presents rich new families of relative orbits for spacecraft formation flight generated through the application of continuous thrust with only minimal intervention into the dynamics of the problem. Such simplicity facilitates implementation for small, low-cost spacecraft with only position state feedback, and yet permits interesting and novel relative orbits in both two- and three-body systems with potential future applications in space-based interferometry, hyperspectral sensing, and on-orbit inspection. Position feedback is used to modify the natural frequencies of the linearised relative dynamics through direct manipulation of the system eigenvalues, producing new families of stable relative orbits. Specifically, in the Hill–Clohessy–Wiltshire frame, simple adaptations of the linearised dynamics are used to produce a circular relative orbit, frequency-modulated out-of-plane motion, and a novel doubly periodic cylindrical relative trajectory for the purposes of on-orbit inspection. Within the circular restricted three-body problem, a similar minimal approach with position feedback is used to generate new families of stable, frequency-modulated relative orbits in the vicinity of a Lagrange point, culminating in the derivation of the gain requirements for synchronisation of the in-plane and out-of-plane frequencies to yield a singly periodic tilted elliptical relative orbit with potential use as a Lunar far-side communications relay. The \documentclass[12pt]{minimal} \usepackage{amsmath} \usepackage{wasysym} \usepackage{amsfonts} \usepackage{amssymb} \usepackage{amsbsy} \usepackage{mathrsfs} \usepackage{upgreek} \setlength{\oddsidemargin}{-69pt} \begin{document}$$\Delta v$$\end{document}Δv requirements for the cylindrical relative orbit and singly periodic Lagrange point orbit are analysed, and it is shown that these requirements are modest and feasible for existing low-thrust propulsion technology.

This paper proposes an attitude tracking control for a rigid spacecraft that adapts to two types of faults that commonly occur in reaction wheels: a gain fault and a deviation fault. In its normal operating mode the tracking controller replicates that of a continuous quaternion feedback controller. When a fault occurs in the system the attitude of the spacecraft will deviate from the reference trajectory and will consequently trigger a sliding mode response of the control which introduces robustness. For the proposed control law, we construct a suitable Lyapunov function to prove the closed-loop system is asymptotically stable in the presence of such faults. However, the proposed control is not practically suitable over long periods as the gain on the sliding mode component will always increase unless the sliding surface is exactly zero (in practise this is never the case because of sensor noise). To address this problem a simple adaptive parameter is defined such that it converges to an appropriate upper-limit. Simulations of the attitude dynamics of a spacecraft are undertaken which compares the tracking performance in the presence of a fault with and without the adaptive sliding mode component.

Attitude slew motions for spacecraft are usually undertaken using feedback control where only the desired final attitude is stated. In this paper attitude guidance is considered which could be used, in addition to feedback control, to enhance the efficiency of slew motions by pre-planning time-dependent attitude motions. This is achieved using a three-step method in which the angular velocities are expressed as analytic functions in terms of free parameters (on the virtual time domain), and the boundary conditions on the rotation are matched using a shooting method based on a discretized form of Rodrigue's formula. Following this, the virtual time is reparametrized. This is applied to design a rest-to-rest two-impulse slew manoeuver and a slew motion using only two reaction wheels.

In this paper the natural dynamics of a rigid body are exploited to plan attitude manoeuvres for a small spacecraft. By utilising the analytical solutions of the angular velocities and making use of Lax pair integration, the time evolution of the attitude of the spacecraft in a convenient quaternion form is derived. This enables repointing manoeuvres to be generated by optimising the free parameters of the analytical expressions, the initial angular velocities of the spacecraft, to match prescribed boundary conditions on the final attitude of the spacecraft. This produces reference motions which can be tracked using a simple proportional-derivative controller. The natural motions are compared in simulation to a conventional quaternion feedback controller and found to require lower accumulated torque. A simple obstacle avoidance algorithm, exploiting the analytic form of natural motions, is also described and implemented in simulation. The computational efficiency of the motion planning method is discussed.

One of the paramount stepping stones towards NASA׳s long-term goal of undertaking human missions to Mars is the exploration of the Martian moons. Since a precursor mission to Phobos would be easier than landing on Mars itself, NASA is targeting this moon for future exploration, and ESA has also announced Phootprint as a candidate Phobos sample-and-return mission. Orbital dynamics around small planetary satellites are particularly complex because many strong perturbations are involved, and the classical circular restricted three-body problem (R3BP) does not provide an accurate approximation to describe the system׳s dynamics. Phobos is a special case, since the combination of a small mass-ratio and length-scale means that the sphere-of-influence of the moon moves very close to its surface. Thus, an accurate nonlinear model of a spacecraft׳s motion in the vicinity of this moon must consider the additional perturbations due to the orbital eccentricity and the complete gravity field of Phobos, which is far from a spherical-shaped body, and it is incorporated into an elliptic R3BP using the gravity harmonics series-expansion (ER3BP-GH). In this paper, a showcase of various classes of non-keplerian orbits is identified and a number of potential mission applications in the Mars–Phobos system are proposed: these results could be exploited in upcoming unmanned missions targeting the exploration of this Martian moon. These applications include: low-thrust hovering and orbits around Phobos for close-range observations; the dynamical substitutes of periodic and quasi-periodic Libration Point Orbits in the ER3BP-GH to enable unique low-cost operations for space missions in the proximity of Phobos; their manifold structure for high-performance landing/take-off maneuvers to and from Phobos’ surface and for transfers from and to Martian orbits; Quasi-Satellite Orbits for long-period station-keeping and maintenance. In particular, these orbits could exploit Phobos’ occulting bulk and shadowing wake as a passive radiation shield during future manned flights to Mars to reduce human exposure to radiation, and the latter orbits can be used as an orbital garage, requiring no orbital maintenance, where a spacecraft could make planned pit-stops during a round-trip mission to Mars.