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Space mission design places a premium on cost and operational efficiency. The search for new science and life beyond Earth calls for spacecraft that can deliver scientific payloads to geologically rich yet hazardous landing sites. At the same time, the last four decades of optimization research have put a suite of powerful optimization tools at the fingertips of the controls engineer. As we enter the new decade, optimization theory, algorithms, and software tooling have reached a critical mass to start seeing serious application in space vehicle guidance and control systems. This survey paper provides a detailed overview of recent advances, successes, and promising directions for optimization-based space vehicle control. The considered applications include planetary landing, rendezvous and proximity operations, small body landing, constrained attitude reorientation, endo-atmospheric flight including ascent and reentry, and orbit transfer and injection. The primary focus is on the last ten years of progress, which have seen a veritable rise in the number of applications using three core technologies: lossless convexification, sequential convex programming, and model predictive control. The reader will come away with a well-rounded understanding of the state-of-the-art in each space vehicle control application, and will be well positioned to tackle important current open problems using convex optimization as a core technology. [ABSTRACT FROM AUTHOR]

This paper addresses the approaching trajectory control problem for the final phase of non-cooperative spacecraft proximity operations in the presence of disturbances and even input saturation. As a stepping stone, a novel line-of-sight tracking model is developed to describe attitude maneuver towards the target in the proximity. Then, a two-stage artificial potential function is employed to enable the pursuer spacecraft to comply the safety trajectory during approaching the target. In each stage, gradients of potential function are analyzed and the local minimum problem associated with potential function can be addressed. Furthermore, a saturated control law is proposed to ensure a reliable real-time tracking measurement of the target, and the controller rigorously enforces safe path and actuator magnitude constraints. The associated stability proof is accomplished through Lyapunov method. Simulations are carried out to evaluate the effectiveness of the proposed method. [ABSTRACT FROM AUTHOR]

Abstract A saturated fault-tolerant attitude tracking controller for disturbed rigid spacecraft is derived using nonlinear state feedback control method. The proposed controller achieves the constraints of control inputs by directly using the bounded function instead of the traditional saturation compensator technique, and the active tolerance to the partial loss of actuator effectiveness is also achieved by directly using the known bounds of the actuator faults in the controller. Specifically, compared with the traditional saturated control methods, a continuously bounded nonlinear function in the proposed controller is used to guarantee that the actuator outputs are smoothly bounded under the prescribed constraints. Based on some properties of the attitude tracking dynamics, the proposed controller can ensure the attitude tracking errors converge to small neighborhoods of zero via stability analysis in the Lyapunov framework. Simulation results are presented to illustrate the effectiveness of the control scheme. [ABSTRACT FROM AUTHOR]

Abstract Motions of spacecraft in non-synchronized binary asteroid systems are facing significant dynamics problems due to the irregular-shaped primaries and time-varying gravitational field. This paper presents a novel method for the design and maintenance of bounded trajectory in the vicinity of non-synchronized binary systems. A non-synchronized model with considering both the planar full two-body problem and the elliptic restricted three-body problem is developed for the bounded motion design. This model can better reflect the properties of motion in non-synchronized binary systems. Based on this non-synchronized model, bounded motions are designed with predetermined stability using the modified grid search and two-level correction strategy. Further, considering the model uncertainty and unmodeled perturbations in the high-fidelity model, a self-adaptive control method is proposed to maintain the trajectories. Simulations for the binary asteroid 1999 KW4 validate the stability and robustness of the proposed controller. Compared with periodic orbits using the synchronized model, the proposed strategy by taking the bounded trajectories as reference trajectories can effectively improve the tracking accuracy and reduce the fuel consumption. This provides a more effective solution for mission trajectory design in future explorations to binary asteroid systems. Highlights • The spacecraft dynamics in the non-synchronized binary system is developed. • The bounded trajectory design for the non-synchronized binary system is proposed. • A self-adaptive control method is studied for orbit maintenance in binary systems. [ABSTRACT FROM AUTHOR]

Modern spacecraft are usually equipped with various flexible appendages, which bring great challenges to the design of high-performance control systems. This paper proposes a novel approach to provide active damping to spacecraft flexible appendages. Vibrations of flexible appendages are actively attenuated using a base-mounted angular displacement mechanism and tip-mounted non-collocated sensors. Dynamic models are presented and the controller design techniques provided. Effectiveness of the approach is demonstrated through an experimental apparatus, in which the damping ratio of the target beam is increased to be better than 40%. Furthermore, an active damping-control scheme using no extra sensors beyond attitude sensors onboard the spacecraft bus is proposed and demonstrated through numerical simulations. [ABSTRACT FROM AUTHOR]

The paper deals with the problem of time-optimal spacecraft reorientation maneuvers by means of reaction wheels, with boundary and path constraints. When searching for solutions to optimal attitude-control problems, spacecraft can be easily modeled as controlled by external torques. However, when using actuators such as reaction wheels, conservation of the total angular momentum must be taken into account and the wheel dynamics must be included. A rest-to-rest slew maneuver is considered where an optical sensor cannot be exposed to sources of bright light such as the Earth, the Sun and the Moon. The motion must be constrained to prevent the sensor axis from entering into established keep-out cones. The minimum-time solution is proposed using the Inverse Dynamics Particle Swarm Optimization technique. The attitude and the kinematics of the satellite evolve, leading to the successive attainment of the wheel control input via fixed-step numerical integration. Numerical results are evaluated over different scenarios. It is established that the computation of minimum time maneuvers with the proposed technique leads to near optimal solutions, which fully satisfy all the boundary and path constraints. The ability to converge in a variety of different scenarios always requiring the same computational effort characterizes the proposed technique as a feasible future on-board path-planner. [ABSTRACT FROM AUTHOR]

Due to configurability feature and increasingly complex architecture, FPGAs have brought advantages to many applications such as avionics and safety critical aerospace, allowing in system reconfiguration after launch. Commercial FPGAs suffer from radiation-induced failures, which are provoked by high-energy particles in space; for this reason, fault tolerant techniques are necessary to harden these devices. This paper presents a design of a fault tolerant multicore processor architecture based on a novel modular voting strategy that fits in FPGAs and System-on-Chip (SoC) devices with an even number of processors. This architecture is implemented within a Commercial off-the-shelf (COTS) SoC that will allow to be used safely in space missions. To harden the fault tolerance of the embedded multicore processor architecture different fault tolerance techniques are combined. [ABSTRACT FROM AUTHOR]

A rotary single inverted pendulum (RSIP) typically represents a space booster rocket, Segway and similar systems with unstable equilibrium. This paper proposes a novel two degree of freedom (2-DOF) fractional control strategy based on 2-loop topology for RSIP system which can be extended to control the systems with unstable equilibrium. It comprises feedback and feed-forward paths. Primary controller relates the perturbation attenuation while the secondary controller is accountable for set point tracking. To tune the parameters of proposed fractional controller a simple graphical tuning method based on frequency response is used. The study will serve the outstanding experimental results for both, stabilization and trajectory tracking tasks. The study will also serve to present a comparison of the performance of the proposed controller with the 1-DOF FOPID controller and sliding mode controller (SMC) for the RSIP system. Further to confirm the usability of the proposed controller and to avoid the random perturbations sensitivity, robustness, and stability analysis through fractional root-locus and Bode-plot is investigated. [ABSTRACT FROM AUTHOR]

A novel star identification algorithm based on voting scheme is presented in this paper. In the proposed algorithm, the global distribution and local distribution of sensor stars are fully utilized, and the stratified voting scheme is adopted to obtain the candidates for sensor stars. The database optimization is employed to reduce its memory requirement and improve the robustness of the proposed algorithm. The simulation shows that the proposed algorithm exhibits 99.81% identification rate with 2-pixel standard deviations of positional noises and 0.322-Mv magnitude noises. Compared with two similar algorithms, the proposed algorithm is more robust towards noise, and the average identification time and required memory is less. Furthermore, the real sky test shows that the proposed algorithm performs well on the real star images. [ABSTRACT FROM AUTHOR]

This paper presents a trajectory-optimization method for a chaser spacecraft operating in ultra-close proximity to a failed satellite. Based on the combination of active and passive trajectory protection, the constraints in the optimization framework are formulated for collision avoidance and successful docking in the presence of any thruster failure. The constraints are then handled by an adaptive Gauss pseudospectral method, in which the dynamic residuals are used as the metric to determine the distribution of collocation points. A finite-time feedback control is further employed in tracking the optimized trajectory. In particular, the stability and convergence of the controller are proved. Numerical results are given to demonstrate the effectiveness of the proposed methods. [ABSTRACT FROM AUTHOR]

The deployment dynamics of a spin stabilized electric sail (E-sail) with a hub-mounted control actuator are investigated. Both radial and tangential deployment mechanisms are considered to take the electric sail from a post-launch stowed configuration to a fully deployed configuration. The tangential configuration assumes the multi-kilometer tethers are wound up on the exterior of the spacecraft hub, similar to yo-yo despinner configurations. The deployment speed is controlled through the hub rate. The radial deployment configuration assumes each tether is on its own spool. Here both the hub and spool rate are control variables. The sensitivity of the deployment behavior to E-sail length, maximum rate and tension parameters is investigated. A constant hub rate deployment is compared to a time varying hub rate that maintains a constant tether tension condition. The deployment time can be reduced by a factor of 2 or more by using a tension controlled deployment configuration. [ABSTRACT FROM AUTHOR]

This paper proposes a novel feedback control law for spacecraft to deal with attitude constraint, input saturation, and stochastic disturbance during the attitude reorientation maneuver process. Applying the parameter selection method to improving the existence conditions for the repulsive potential function, the universality of the potential-function-based algorithm is enhanced. Moreover, utilizing the auxiliary system driven by the difference between saturated torque and command torque, a backstepping control law, which satisfies the input saturation constraint and guarantees the spacecraft stability, is presented. Unlike some methods that passively rely on the inherent characteristic of the existing controller to stabilize the adverse effects of external stochastic disturbance, this paper puts forward a nonlinear disturbance observer to compensate the disturbance in real-time, which achieves a better performance of robustness. The simulation results validate the effectiveness, reliability, and universality of the proposed control law. [ABSTRACT FROM AUTHOR]

Observability analysis is a method for determining whether a chosen state of a system can be determined from the output or measurements. Knowledge of state information availability resulting from observability analysis leads to improved sensor tasking for observation of orbital debris and better control of active spacecraft. This research performs numerical observability analysis of artificial near-Earth objects. Analysis of linearization methods and state transition matrices is performed to determine the viability of applying linear observability methods to the nonlinear orbit problem. Furthermore, pre-whitening is implemented to reformulate classical observability analysis. In addition, the state in observability analysis is typically composed of position and velocity; however, including object characteristics beyond position and velocity can be crucial for precise orbit propagation. For example, solar radiation pressure has a significant impact on the orbit of high area-to-mass ratio objects in geosynchronous orbit. Therefore, determining the time required for solar radiation pressure parameters to become observable is important for understanding debris objects. In order to compare observability analysis results with and without measurement noise and an extended state, quantitative measures of observability are investigated and implemented. [ABSTRACT FROM AUTHOR]

In this paper, the attitude stabilization and tracking control problem is investigated for spacecraft with consideration of unknown dynamics and external disturbance. A low-complexity prescribed performance attitude control scheme is presented to solve this problem. Different from the existing works, there are threefold prominent advantages. The first one is the transient performance and the steady performance of the system is guaranteed by a user-defined function rather than depending on repeated adjustment of controller parameters. The second is that no information of the system and external disturbance is necessary in the developed control scheme, which means the method is model-free. Moreover, the developed low-complexity controller is calculated without any time-consuming iterative operations; thus it's significantly advantageous in engineering applications. It is proved that the state variables converge to the prescribed region at a prescribed exponential rate under the proposed control scheme. Four groups of numerical simulations are organized to validate the effectiveness of the method. [ABSTRACT FROM AUTHOR]

In this paper, a robust inertia-free attitude takeover control scheme with guaranteed prescribed performance is investigated for postcapture combined spacecraft with consideration of unmeasurable states, unknown inertial property and external disturbance torque. Firstly, to estimate the unavailable angular velocity of combination accurately, a novel finite-time-convergent tracking differentiator is developed with a quite computationally achievable structure free from the unknown nonlinear dynamics of combined spacecraft. Then, a robust inertia-free prescribed performance control scheme is proposed, wherein, the transient and steady-state performance of combined spacecraft is first quantitatively studied by stabilizing the filtered attitude tracking errors. Compared with the existing works, the prominent advantage is that no parameter identifications and no neural or fuzzy nonlinear approximations are needed, which decreases the complexity of robust controller design dramatically. Moreover, the prescribed performance of combined spacecraft is guaranteed a priori without resorting to repeated regulations of the controller parameters. Finally, four illustrative examples are employed to validate the effectiveness of the proposed control scheme and tracking differentiator. [ABSTRACT FROM AUTHOR]

This paper presents an adaptive neural networks-based control method for spacecraft formation with coupled translational and rotational dynamics using only aerodynamic forces. It is assumed that each spacecraft is equipped with several large flat plates. A coupled orbit-attitude dynamic model is considered based on the specific configuration of atmospheric-based actuators. For this model, a neural network-based adaptive sliding mode controller is implemented, accounting for system uncertainties and external perturbations. To avoid invalidation of the neural networks destroying stability of the system, a switching control strategy is proposed which combines an adaptive neural networks controller dominating in its active region and an adaptive sliding mode controller outside the neural active region. An optimal process is developed to determine the control commands for the plates system. The stability of the closed-loop system is proved by a Lyapunov-based method. Comparative results through numerical simulations illustrate the effectiveness of executing attitude control while maintaining the relative motion, and higher control accuracy can be achieved by using the proposed neural-based switching control scheme than using only adaptive sliding mode controller. [ABSTRACT FROM AUTHOR]

High radio-frequency space applications have motivated continuous research on vibration suppression of large space structures both in academia and industry. This paper introduces a novel data driven control method to suppress vibrations of flexible structures and experimentally validates the suppression performance. Unlike model-based control approaches, the data driven control method designs a controller directly from the input-output test data of the structure, without requiring parametric dynamics and hence free of system modeling. It utilizes the discrete frequency response via spectral analysis technique and formulates a non-convex optimization problem to obtain optimized controller parameters with a predefined controller structure. Such approach is then experimentally applied on an end-driving flexible beam-mass structure. The experiment results show that the presented method can achieve competitive disturbance rejections compared to a model-based mixed sensitivity controller under the same design criterion but with much less orders and design efforts, demonstrating the proposed data driven control is an effective approach for vibration suppression of flexible structures. [ABSTRACT FROM AUTHOR]

The stability characteristics of a distributed consensus-based Kalman filter estimation and control scheme are studied through analytic and numerical means. This estimation scheme seeks to minimally reduce the necessary bandwidth for communication while maintaining overall stability. A weaker form of the separation principle is proven to hold whereby control could be designed independently but not estimation. However, actuation limitations still provide the possibility for a semi-independent design of estimators. Numerical simulations confirm that the stability depends very heavily on consensus on estimation and ultimately the amount of information available to the system as it evolves. [ABSTRACT FROM AUTHOR]

In this paper, a novel cooperative guidance law is proposed to make multiple missiles in the three-dimensional (3-D) space hit simultaneously the same target at pre-specified impact angles. Firstly, the normal accelerations which change the velocity direction (flight-path and heading angle) are designed such that all missiles will fly along the desired line of sight (LOS) after a given time which ensures the hit-to-kill interception at the desired impact angles; then the consensus variable is constructed using available information and can reach consensus under the proposed tangential acceleration which determines the velocity magnitude. Hence simultaneous hit-to-kill attack is achieved. Finally, some simulation studies are performed to verify the effectiveness of the proposed scheme. [ABSTRACT FROM AUTHOR]

This article investigates the velocity-free attitude coordinated tracking control scheme for a group of spacecraft with the assumption that the angular velocities of the formation members are not available in control feedback. Initially, an angular velocity observer is constructed based on each individual's attitude quarternion. Then, the distributed attitude coordinated control law is designed by using the observed states, in which adaptive control method is adopted to handle the external disturbances. Stability of the overall closed-loop system is analyzed theoretically, which shows the system trajectory converges to a small set around origin with fast convergence rate. Numerical simulations are performed to demonstrate fast convergence and improved tracking performance of the proposed control strategy. [ABSTRACT FROM AUTHOR]

Based on Nodal Coordinate Formulation (NCF) and Absolute Nodal Coordinate Formulation (ANCF), this paper establishes rigid-flexible coupling dynamic model of the spacecraft with large deployable solar arrays and multiple clearance joints to analyze and control the satellite attitude under deployment disturbance. Considering torque spring, close cable loop (CCL) configuration and latch mechanisms, a typical spacecraft composed of a rigid main-body described by NCF and two flexible panels described by ANCF is used as a demonstration case. Nonlinear contact force model and modified Coulomb friction model are selected to establish normal contact force and tangential friction model, respectively. Generalized elastic force are derived and all generalized forces are defined in the NCF-ANCF frame. The Newmark- β method is used to solve system equations of motion. The availability and superiority of the proposed model is verified through comparing with numerical co-simulations of Patran and ADAMS software. The numerical results reveal the effects of panel flexibility, joint clearance and their coupling on satellite attitude. The effects of clearance number, clearance size and clearance stiffness on satellite attitude are investigated. Furthermore, a proportional-differential (PD) attitude controller of spacecraft is designed to discuss the effect of attitude control on the dynamic responses of the whole system. [ABSTRACT FROM AUTHOR]

Tethered space-tug system is regarded as one of the most promising active debris removal technologies to effectively decrease the steep increasing population of space debris. In order to suppress the spin of space debris, single-tethered space-tug system is employed by regulating the tether. Unfortunately, this system is underactuated as tether length is the only input, and there are two control objectives: the spinning debris and the vibration of tether. Thus, it may suffer great oscillations and result in failure in space debris removal. This paper presents the study of attitude stabilization of the single-tethered space-tug system using not only tether length but also the offset of tether attachment point to suppress the spin of debris, so as to accomplish the space debris removal mission. Firstly, a precise 3D mathematical model in which the debris and tug are both treated as rigid bodies is developed to study the dynamical evolution of the tethered space-tug system. The relative motion equation of the system is described using Lagrange method. Secondly, the dynamic characteristic of the system is analyzed and an offset control law is designed to stabilize the spin of debris by exploiting the variation of tether offset and the regulation of tether length. Besides, an estimation formula is proposed to evaluate the capability of tether for suppressing spinning debris. Finally, the effectiveness of attitude stabilization by the utilization of the proposed scheme is demonstrated via numerical case studies. [ABSTRACT FROM AUTHOR]

Two-phase micro-channel heat sinks are prime candidates for incorporation into thermal control systems (TCSs) of future space vehicles and planetary habitats. Unlike small heat sinks employed in the electronics industry, TCS heat sinks are characterized by large length-to-diameter ratio, for which limited information is presently available. This study employs a 609.6-mm long by 203.2-mm wide heat sink containing 100 of 1 × 1 mm 2 micro-channels and uses R134a as working fluid. The large length-to-diameter ratio of 609.6 is especially instrumental to capturing detailed axial variations of flow pattern and corresponding variations in local heat transfer coefficient. High-speed video analysis of the inlet plenum shows appreciable vapor backflow under certain operating conditions, which is also reflected in periodic oscillations in the measured pressure drop. In fact, the backflow frequency captured by video matches closely the frequency obtained from Fourier analysis of the pressure drop signal. While density-wave oscillations are encountered in individual channels, the phenomena observed are more closely related to parallel-channel instability. It is shown the periodic oscillations and vapor backflow are responsible for initiating intermittent dryout and appreciable drop in local heat transfer coefficient in the downstream regions of the channels. A parametric study of oscillation frequency shows a dependence on four dimensionless parameters that account for amount of vapor generation, subcooling, and upstream liquid length, in addition to Weber number. A new correlation for oscillation frequency is constructed that captures the frequency variations relative to these individual parameters. [ABSTRACT FROM AUTHOR]

Finite-time controllers are proposed in this paper for underactuated spacecraft hovering in the absence of the radial or in-track thrust. The indirect method, which is generally adopted to solve the singularity problem in the conventional terminal sliding mode, is modified to ensure the continuity of the high-order time derivative of the sliding surface at the switch points. Rigorous proofs via the Lyapunov-based approaches verify the finite-time stability of the closed-loop system. By comparisons with the asymptotic controllers, the advantages of the finite-time ones in faster convergence rate and enhanced control precision have also been substantiated. [ABSTRACT FROM AUTHOR]

In this paper, a novel approach for high-accuracy attitude control in a flexible spacecraft is developed by compensating and damping the vibration of the flexible appendage. A vibration suppression-based attitude controller (VSAC) that addresses the coupling disturbance in attitude dynamics that is generated by the vibration of the flexible appendage is developed for flexible spacecraft. The proposed VSAC integrates the existing attitude control theories with vibration suppression control strategies by combining a time delay feedback controller with a PD controller. The stability of the flexible spacecraft control system under the proposed VSAC is investigated in time domain, and the damping effect of the proposed VSAC on the vibration of the flexible appendage is analyzed in frequency domain. The proposed VSAC utilizes the traditional attitude controller and vibration suppression controller, which can not only effectively improve the performance of attitude control system but also efficiently damp the vibration of the flexible appendage. Finally, numerical simulations are performed, and analysis and conclusions are presented. [ABSTRACT FROM AUTHOR]

We analyze a cooperative control framework for electric sail formation flying around a heliocentric displaced orbit, aiming at observing the polar region of a celestial body. The chief spacecraft is assumed to move along an elliptic displaced orbit, while each deputy spacecraft adjusts its thrust vector (that is, both its sail attitude and characteristic acceleration) in order to track a prescribed relative trajectory. The relative motion of the electric sail formation system is formulated in the chief rotating frame, where the control inputs of each deputy are the relative sail attitude angles and the relative lightness number with respect to those of the chief. The information exchange among the spacecraft, characterized by the communication topology, is represented by a weighted graph. Two typical cases, according to whether the communication graph is directed or undirected, are discussed. For each case, a distributed coordinated control law is designed in such a way that each deputy not only tracks the chief state, but also makes full use of information from its neighbors, thus increasing the redundancy and robustness of the formation system in case of failure among the communication links. Illustrative examples show the effectiveness of the proposed approach. [ABSTRACT FROM AUTHOR]

Deployment systems for innovative space applications such as solar sails require a technique for a controlled and autonomous deployment in space. The deployment process has a strong impact on the mechanism and structural design and sizing. On the example of the design implemented in the Gossamer-1 project of the German Aerospace Center (DLR), such a stowing and deployment process is analyzed. It is based on a combination of zig-zag folding and coiling of triangular sail segments spanned between crossed booms. The deployment geometry and forces introduced by the mechanism considered are explored in order to reveal how the loads are transferred through the membranes to structural components such as the booms. The folding geometry and force progressions are described by function compositions of an inverse trigonometric function with the considered trigonometric function itself. If these functions are evaluated over several periods of the trigonometric function, a non-smooth oscillating curve occurs. Depending on the trigonometric function, these are often vividly described as zig-zag or sawtooth functions. The developed functions are applied to the Gossamer-1 design. The deployment geometry reveals a tendency that the loads are transferred along the catheti of the sail segments and therefore mainly along the boom axes. The load introduced by the spool deployment mechanism is described. By combining the deployment geometry with that load, a prediction of the deployment load progression is achieved. The mathematical description of the stowing and deployment geometry, as well as the forces inflicted by the mechanism provides an understanding of how exactly the membrane deploys and through which edges the deployment forces are transferred. The mathematical analysis also gives an impression of sensitive parameters that could be influenced by manufacturing tolerances or unsymmetrical deployment of the sail segments. While the mathematical model was applied on the design of the Gossamer-1 hardware, it allows an analysis of other geometries. This is of particular interest as Gossamer-1 investigated deployment technology on a relatively small scale of 5 m × 5 m , while the currently considered solar sail missions require sails that are about one order of magnitude bigger. [ABSTRACT FROM AUTHOR]

The main problem addressed is the quaternion-based attitude stabilization control of rigid spacecraft without angular velocity measurements in the presence of external disturbances and reaction wheel friction as well. As a stepping stone, an angular velocity observer is proposed for the attitude control of a rigid body in the absence of angular velocity measurements. The observer design ensures finite-time convergence of angular velocity state estimation errors irrespective of the control torque or the initial attitude state of the spacecraft. Then, a novel finite-time control law is employed as the controller in which the estimate of the angular velocity is used directly. It is then shown that the observer and the controlled system form a cascaded structure, which allows the application of the finite-time stability theory of cascaded systems to prove the finite-time stability of the closed-loop system. A rigorous analysis of the proposed formulation is provided and numerical simulation studies are presented to help illustrate the effectiveness of the angular-velocity observer for rigid spacecraft attitude control. [ABSTRACT FROM AUTHOR]

This study addresses a relative position coordinated control problem for spacecraft formation flying subject to directed communication topology. Two different kinds of communication delay cases, including time-varying delays and arbitrarily bounded delays are investigated. Using the backstepping control technique, two virtual velocity control inputs are firstly designed to achieve coordinated position tracking for the kinematic subsystem. Furthermore, a hyperbolic tangent function is introduced to guarantee the boundedness of the virtual controller. Then, a finite-time control algorithm is designed for the dynamic subsystem. It can guarantee that the virtual velocity can be followed by the real velocity after finite time. It is theoretically proved that the proposed control scheme can asymptotically stabilize the closed-loop system. Numerical simulations are further presented that not only highlight closed-loop performance benefiting from the proposed control scheme, but also illustrate its superiority in comparison with conventional formation control schemes. [ABSTRACT FROM AUTHOR]

This paper investigates the finite time coordinated formation control problem for spacecraft formation flying (SFF) under the assumption of directed communication topology. By using the neighborhood state measurements, a robust finite time coordinated formation controller is firstly designed based on the nonsingular terminal sliding mode surface. To address the special case that the desired trajectory of the formation is only accessible to a subset of spacecraft in the formation, an adaptive finite time coordinated formation controller is also proposed by designing a novel sliding mode surface. In both cases, the external disturbances are explicitly taken into account. Rigorous theoretical analysis proves that the proposed control schemes ensure that the closed-loop system can track the desired time-varying trajectory in finite time. Numerical simulations are presented that not only highlights the closed-loop performance benefits from the proposed control algorithms, but also illustrates the effectiveness in the presence of external disturbances when compared with the existing coordinated formation control schemes. [ABSTRACT FROM AUTHOR]

A method for asteroid deflection that makes use of a spacecraft moving back and forth on a segment of a Keplerian orbit about the asteroid is described and studied. It is shown that on average the spacecraft can exert a significantly larger force on the asteroid than e.g. a stationary gravity tractor, thereby reducing the time needed to effect the desired deflection of the asteroid. Furthermore, the current method does not require canted thrusters on the spacecraft (unlike a stationary gravity tractor) markedly reducing the amount of fuel needed for a given deflection to be realized. The method also allows for the simultaneous use of several spacecraft, further strengthening the overall tugging effect on the asteroid, and distributing the thrust requirement among the spacecraft. [ABSTRACT FROM AUTHOR]

Up to April 6, 2016, there are 17,385 large debris in orbit around the Earth, which poses a serious hazard to near-Earth space activities. As a promising on-orbit debris capture strategy, tethered space robots (TSRs) have wide applications in future on-orbit service owing to its flexibility and great workspace. However, lots of problems may arise in the Tethered Space Robots (TSRs) system from the approaching, capturing, postcapturing and towing phases. The postcapture combination attitude takeover control by the TSR is studied in this paper. Taking control constraints, tether oscillations and external disturbances into consideration, a fast terminal sliding mode control (FTSMC) methodology with dual closed loops for the flexible combination attitude takeover control is designed. The unknown upper bounds of the uncertainties, external disturbances are estimated through adaptive techniques. Stability of the dual closed loop control system and finite time convergence of system states are proved via Lyapunov stability theory. Besides, null space intersection control allocation was adopted to distribute the required control moment over TSR's redundant thrusters. Simulation studies have been conducted to demonstrate the effectiveness of the proposed controller with the conventional sliding mode control(SMC). [ABSTRACT FROM AUTHOR]

Advanced navigation systems for pinpoint landing are required in the entry, descent and landing phase of future missions to Mars. To overcome the horizontal position estimation problem in the Mars powered descent phase, the inertial measurement unit, Doppler radar and surface beacon integrated navigation scheme is proposed, and a conventional filter such as an extended or unscented Kalman filter is adopted. However, in engineering practice, these conventional filters for nonlinear systems with unknown dynamic inputs may degrade or even diverge. The lack of navigation accuracy may result in a large growth of spurious navigations. To solve this problem, based on the rank filter, a self-calibration rank filter is proposed for state estimation of a nonlinear system to mitigate the effects of unknown dynamic inputs. Monte Carlo simulation results are presented to demonstrate the good performance of the self-calibration rank filter for the Mars powered descent navigation. The self-calibration rank filter not only prevents the divergence of the filtering but also significantly improves the state estimation accuracy. [ABSTRACT FROM AUTHOR]

Areostationary satellites are considered a high interest group of satellites to satisfy the telecommunications needs of the foreseen missions to Mars. An areostationary satellite, in an areoequatorial circular orbit with a period of 1 Martian sidereal day, would orbit Mars remaining at a fixed location over the Martian surface, analogous to a geostationary satellite around the Earth. This work addresses an analysis of the perturbed orbital motion of an areostationary satellite as well as a preliminary analysis of the aerostationary orbit estimation accuracy based on Earth tracking observations. First, the models for the perturbations due to the Mars gravitational field, the gravitational attraction of the Sun and the Martian moons, Phobos and Deimos, and solar radiation pressure are described. Then, the observability from Earth including possible occultations by Mars of an areostationary satellite in a perturbed areosynchronous motion is analyzed. The results show that continuous Earth-based tracking is achievable using observations from the three NASA Deep Space Network Complexes in Madrid, Goldstone and Canberra in an occultation-free scenario. Finally, an analysis of the orbit determination accuracy is addressed considering several scenarios including discontinuous tracking schedules for different epochs and different areoestationary satellites. Simulations also allow to quantify the aerostationary orbit estimation accuracy for various tracking series durations and observed orbit arc-lengths. [ABSTRACT FROM AUTHOR]

The minimum-fuel orbital transfer is analyzed in the case of a launcher upper stage using a constantly thrusting engine. The thrust level is assumed to be constant and its value is optimized together with the thrust direction. A closed-loop solution for the thrust direction is derived from the extremal analysis for a planar orbital transfer. The optimal control problem reduces to two unknowns, namely the thrust level and the final time. Guessing and propagating the costates is no longer necessary and the optimal trajectory is easily found from a rough initialization. On the other hand the initial costates are assessed analytically from the initial conditions and they can be used as initial guess for transfers at different thrust levels. The method is exemplified on a launcher upper stage targeting a geostationary transfer orbit. [ABSTRACT FROM AUTHOR]

Jettisoned Entry, Descent and Landing System (EDLS) hardware from landing spacecraft have been observed by orbiting spacecraft, strewn over the Martian surface. Future Mars missions that land spacecraft close to prelanded assets will have to use a landing architecture that somehow minimises the possibility of impacts from these jettisoned EDLS components. Computer modelling is used here to investigate the influence of wind speed and direction on the distribution of EDLS components on the surface. Typical wind speeds encountered in the Martian Planetary Boundary Layer (PBL) were found to be of sufficient strength to blow items having a low ballistic coefficient, i.e. Hypersonic Inflatable Aerodynamic Decelerators (HIADs) or parachutes, onto prelanded assets even when the lander itself touches down several kilometres away. Employing meteorological measurements and careful characterisation of the Martian PBL, e.g. appropriate wind speed probability density functions, may then benefit future spacecraft landings, increase safety and possibly help reduce the delta v budget for Mars landers that rely on aerodynamic decelerators. [ABSTRACT FROM AUTHOR]

A novel improved mixed H 2 / H ∞ control technique combined with poles assignment theory is presented to achieve attitude stabilization and vibration suppression simultaneously for flexible spacecraft in this paper. The flexible spacecraft dynamics system is described and transformed into corresponding state space form. Based on linear matrix inequalities (LMIs) scheme and poles assignment theory, the improved mixed H 2 / H ∞ controller does not restrict the equivalence of the two Lyapunov variables involved in H 2 and H ∞ performance, which can reduce conservatives compared with traditional mixed H 2 / H ∞ controller. Moreover, it can eliminate the coupling of Lyapunov matrix variables and system matrices by introducing slack variable that provides additional degree of freedom. Several simulations are performed to demonstrate the effectiveness and feasibility of the proposed method in this paper. [ABSTRACT FROM AUTHOR]

The problem of spacecraft attitude tracking using hybrid actuators with uncertainties is addressed in this paper. A hybrid actuators configuration that combines reaction wheels for fine pointing and single gimbal control moment gyros for rapid maneuvering is employed for agile spacecraft. A robust control algorithm for the spacecraft attitude tracking problem when the torque axis direction and/or input scaling of the actuators are uncertain is developed. Furthermore, a torque allocation method is proposed for the hybrid actuator configuration to allow a smooth switch between single gimbal control moment gyros and reaction wheels. With this method, single gimbal control moment gyros are used for the phase of rapid maneuvering, while reaction wheels are used for the phase of fine pointing. Simulation results demonstrate the effectiveness of the proposed control scheme. [ABSTRACT FROM AUTHOR]

A growing interest within the space community focuses on robotics due to the large number of possible applications in many mission scenarios. On-Orbit Servicing (OOS) is arguably the most appealing implementation of space automatic systems. In several cases, OOS requires the capture of orbital objects, which is a complex and risky operation that can be successfully performed by robotic manipulators. Soft robotics, in particular, seems to be suitable for such applications given its intrinsic compliance to the operative environment. Devices based on Dielectric Elastomers (DE) can be employed for the implementation of soft robotic systems and showed promising performances. The introduction of DEs to orbital systems would represent a breakthrough in space technologies. In addition, space conditions could further advantage DE robotics, given the reduced environmental loads experienced and the longer times for operations. Nevertheless, Dielectric Elastomer Actuators (DEA) are a low-TRL (Technology Readiness Level) technology that needs to prove its maturity and suitability to space implementation. In this work, the performances of a redundant manipulator based on DEAs are presented in terms of numerical and experimental results. A 4-DoF planar manipulator has been tested in a gravity-compensated setup. The system is composed by two double-cone actuators mounted in series, each of them providing actuation of two DoF. The end-effector is an optical marker whose position is detected by a vision system. The system has a total of four joint DoF and operates in the xy horizontal plane; only the x and y positions of the end-effector are controlled. Two degrees of redundancy are obtained and exploited for the optimization of joint torques to avoid the saturation of actuators. Numerical simulations have been conducted to predict the system behaviour. The laboratory facility emulates the zero-gravity orbital environment by means of a suspending cable. Detailed experimental results are presented and exploited for the validation of control algorithm and numerical models. [ABSTRACT FROM AUTHOR]

This paper presents the design, modeling, control, and experimental results for a one-axis magnetically suspended reaction sphere (1D-MSRS) driven by a hysteresis motor. The goal of this work is twofold: (a) to conduct a preliminary study for magnetically suspended reaction sphere for three-axis spacecraft attitude control, and (b) study the potential of hysteresis motors for the reaction wheel/sphere drives. The 1D-MSRS uses a hysteresis motor with a spherical rotor made of solid steel. The rotor sphere is magnetically suspended in all translational directions, and is driven about the vertical axis by a bearingless hysteresis motor. We present the modeling and control of the magnetic suspension of the bearingless motor in the 1D-MSRS, and the hysteresis motor dynamics are analyzed by a hysteresis motor equivalent circuit model. The 1D-MSRS system has experimentally demonstrated a starting torque of 8.9 mNm under 0.7 A peak sinusoidal excitation current. With this excitation the sphere can run up to 12,000 rpm synchronously in the presence of air drag. This study demonstrates that the hysteresis motor has strong potential for use in high-speed, low-vibration reaction wheels. [ABSTRACT FROM AUTHOR]

This article discusses the design-ballistic optimization for spacecraft, which move under the influence of the gravitational fields of complex configuration and electric propulsion. Such conditions exist near asteroids, for example in the systems of two gravitating bodies, in the vicinity of the libration points. In this paper, we propose to use a method of iterative mission optimization for electric propulsion spacecraft, using sequences of the mathematically specified models of movement and the spacecraft design. The previously obtained analytical results, which describes the planar movement of spacecraft with solar electric propulsion, allows the initial approach in the iterative scheme of optimization to be constructed. A method of modeling and optimization of interplanetary missions ballistic schemes is developed, based on a combination of Pontryagin's maximum principle formalism conditions of transversality and methods of mathematical programming, allowing restriction characteristics for real interplanetary missions to be considered. Recommendations for the choice of the design-ballistic parameters of the interplanetary mission's spacecraft are obtained taking into account the features of a nuclear and solar power plant for purposes of the delivery of a payload to the required heliocentric orbit, to the Moon and near asteroid. [ABSTRACT FROM AUTHOR]

The problem of the optimal control of spacecraft (SC) with an electric ramjet (ER) to change the orbital parameters is considered. ER uses an air as a fuel. Optimal control is determined by solving the problem on the conditional maximum of the local (at the current point of the orbit) impact of the perturbing acceleration on the functional, assuming the perturbations’ smallness compared to gravity. An approximate synthesis of the optimal thrust vector is obtained in view of the power limitation. Parametric investigations of the apogee's changes, depending on the SC with ER parameters, were carried out. The efficiency of the energy and/or fuel accumulation for the implementation of the given orbital flight of SC with ER is shown. [ABSTRACT FROM AUTHOR]

Models of spacecraft guided motion in the gravitational fields of different configurations were developed. Parameterization of the optimal control laws for all the trajectory segments of the spacecraft moving in the N-body gravitational field was realized. The dependence patterns of control laws on mission purposes were determined. The simulation of the number of missions in the Solar system was performed: low Earth orbit – L1 and L2 of the Earth-Moon and the Sun-Earth systems, L1 – L2 and L2 – L1 of the Earth-Moon and the Sun-Earth systems, low Earth orbit – asteroid Eros, low Earth orbit – Venus, low Earth orbit – Mars. Moreover, the Venus gravity assist maneuver was considered. [ABSTRACT FROM AUTHOR]

The problem of cooperative optimization of trajectory and main design parameters of spacecraft, equipped with propulsion systems with energy accumulation, during low Earth orbit formation and correction in non-central gravitational field is considered. The working principle of these propulsion systems is a periodic cycle of energy accumulation from the energy source at the passive stages of orbital transfer, and its discharge to supply power to the engine during the active stages of orbital transfer. The special characteristics of propulsion systems with energy accumulation is limited operation time per one switching, which determines by the energy, stored in an accumulator. For design-ballistic analysis, the mass model of the spacecraft with an uncontrollable power-limited engine with energy accumulator is used. The general problem is separated into dynamic and parametric parts. Using averaging mathematical models of motion and Pontriagin's maximum principle, the calculation method of the ballistic characteristics of approximately-optimal orbital transfer is obtained. The method of design-ballistic optimization is obtained as an iterative procedure. It was found, that inclusion of an energy accumulator in a propulsion system provides an increasing of orbital transfer efficiency. [ABSTRACT FROM AUTHOR]

In the mid-90s of the past century Arnon Spitzer offered a flight scheme comprising an elliptical synchronous equatorial orbit for the SC injection into geostationary orbit (GEO). The advantages of such scheme were related to the SC control simplicity during its transfer from the elliptical synchronous orbit into GEO. The main advantage of Spitzer's scheme is the possibility of observing spacecraft from Earth using a limited number of ground stations along the entire trajectory of the SC flight from an intermediate orbit into GEO. Some flight schemes, which preserve the main advantage of the Spitzer's scheme, are analyzed. The inclination and the eccentricity of the intermediate synchronous orbit are optimized. Several variants of the yaw angle control are analyzed. Numerical analysis was carried out for a space transportation system based on the launch vehicle “Angara-A5”, chemical upper stage “KVTK”, and electric propulsion system comprising four thrusters SPD-140 with the specific impulse of 1700 s. [ABSTRACT FROM AUTHOR]

The paper presents the results of verification of ground testing of the AIST-type small spacecraft thermal control system (TCS) based on telemetry data received during their in-orbit testing and operation by the ground control station (GCS) of Samara University. The study into the efficiency and operability of the TCS was carried out on AIST series small spacecrafts AIST#1 (RS-43as), launched on April 19, 2013 from Baykonur Cosmodrome as a by-cargo of Bion-M #1 spacecraft, and an experimental sample of AIST #2 (RS-41at) small spacecraft launched by Soyuz-2.1 v with the Volga insertion module on December 28, 2013, from the Plesetsk Cosmodrome. [ABSTRACT FROM AUTHOR]

This paper presents the results of verification of ground testing of the AIST-type small spacecraft thermal control system (TCS) based on telemetry data received during their in-orbit testing and operation by the ground control station (GCS) of Samara University. The study into the efficiency and operability of the TCS was done on the AIST series of small spacecrafts AIST#1 (RS-43as), launched on April 19, 2013 from Baykonur Cosmodrome as a by-cargo of Bion-M #1 spacecraft, and an experimental sample of AIST #2 (RS-41at) small spacecraft launched by Soyuz-2.1 v with Volga insertion module on December 28, 2013, from the Plesetsk Cosmodrome. [ABSTRACT FROM AUTHOR]

The problem of controlled spacecraft motion in vicinity of asteroids is an important flight dynamics task. Solution of the problem consists of gravitational field simulations and control trajectory scheme defininition. We propose to use the model of single gravity points for motion simulation of a spacecraft in an irregular gravitational field of the Eros 433 asteroid. The equations of spacecraft motion are the corresponding equations of the n-body problem. The motion trajectory can be split into a number of from-point-to-point motion parts. Each part can be considered as a two-point Boundary task. The Authors developed a concept of Boundary problem solution based on the Pontryagin's principle and modified Newton's step-by-step approximation method. The Authors developed a special software for the Boundary problem solutions. A boundary task of the controlled spacecraft's transfer between circular orbits from 200 km to 100 km is presented in the paper. [ABSTRACT FROM AUTHOR]

The ballistic optimization of the L1-L2 and L2-L1 missions in the Earth-Moon system using electric propulsion is considered. The motion control is studied using the Fedorenko's successful linearization method to estimate the derivatives and the gradient method to optimize the control laws. The transfers are designed in frames of the restricted circular three-body problem. The mathematical model of the missions is described within the barycentric system of coordinates. The perturbation from the Earth, the Moon and the Sun are taking into account. The impact factor of the shaded areas, produced by the Earth and the Moon, is also accounted for. As for the results, the optimization of the optimal control laws and corresponding trajectories were obtained. [ABSTRACT FROM AUTHOR]

The problem of fragment of space debris located at geostationary orbit disposal is studied. For this purpose, a special spacecraft debris collector is used. The main part of disposal operation – rendezvous transfer with fragment of space debris - is studied. Relative motion of a spacecraft debris collector about the fragment of space debris is considered in an orbital cylindrical reference frame. Secular and periodic components of relative motion are determined in an implicit form. Spacecraft debris collector motion control is carried out by its transversal acceleration from thrust switching. Radial thrust component is equal to zero. Boundary conditions for rendezvous transfer are formulated in terms of secular and periodic components of motion. The Motion control program for secular components contained two active areas and passive areas between them. Active areas duration were determined as functions of boundary conditions and passive area duration (free parameter). An analytical solution for the determined control program is formulated in an implicit form. Satisfaction of boundary conditions for periodic components is carried out by passive area duration and rendezvous start moment determination. Equations to determine these parameters are formulated. Numerical simulation is carried out. [ABSTRACT FROM AUTHOR]