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In this paper, the attitude control of spacecraft using a pyramid-type variable-speed control moment gyro (VSCMG) system is considered. For successive attitude maneuvers, it is desirable that the wheel angular momentums are regulated to nominal values after an attitude maneuver. In addition, avoidance of reference aligned singularity (RAS) is also important. RAS is a singularity point where the singular direction of the constant-speed control moment gyro (CSCMG) system is aligned to a reference torque. Regarding the motivations, this paper proposes a steering law which achieves attitude control and regulates the wheel angular momentums while successfully avoiding RAS. Both numerical simulations and experimental verification are performed to demonstrate not only the validity of the proposed method, but also a certain robustness against unmodeled dynamics and the gravity torque. • Spacecraft attitude is controlled by variable speed control moment gyros. • Reference aligned singularities of CMGs are particularly focused in the control. • Wheel angular momentums are usually regulated to nominal values. • Effectiveness of the control law is validated in all maneuver directions. • Ground experiments are conducted in order to examine the proposed control. [ABSTRACT FROM AUTHOR]

This paper focuses on the complex dynamics and control of deploying a payload using a long flexible tether from a spacecraft. Two different system models are involved, one for detailed dynamics study and the other for control design. Both the flexibility and extensibility of tether and the attitude motion of spacecraft have been incorporated into the system dynamics. In addition, the orbital dynamics of the tethered system are modeled using a set of modified equinoctial elements and considering the J 2 perturbation forces. A two-stage mission design has been proposed to divide the deployment into two parts. The first is an ejection separation stage for the payload to be ejected away from the spacecraft to a predetermined position, and the second is a deployment control stage for active tether deployment by regulating tether tension. The attitude motion control of the spacecraft is performed during deployment based on an optimal fuel control method. A closed-loop feedback controller is developed to achieve the deployment by tracking an open-loop optimal trajectory of tether libration states and control input. Finally, numerical simulations have been performed to demonstrate the feasibility of conceptual mission design and effectiveness of the synthetic control scheme. • A two-stage mission design has been proposed. • Spacecraft attitude, tether flexibility and extensibility are accounted for. • Attitude stabilization and tether deployment are simultaneously achieved. [ABSTRACT FROM AUTHOR]

The paper describes the development and simulation results of a full ADCS subsystem for the deOrbitSail drag sail mission. The deOrbitSail satellite was developed as part of an European FP7 collaboration research project. The satellite was launched and commissioning started on 10th July 2015. Various new actuators and sensors designed for this mission will be presented. The deOrbitSail satellite is a 3U CubeSat to deploy a 4 by 4 m drag sail from an initial 650 km circular polar low earth orbit. With an active attitude control system it will be shown that by maximising the drag force, the expected de-orbiting period from the initial altitude will be less than 50 days. A future application of this technology will be the use of small drag sails as low-cost devices to de-orbit LEO satellites, when they have reached their end of life, without having to use expensive propulsion systems. Simulation and Hardware-in-Loop experiments proved the feasibility of the proposed attitude control system. A magnetic-only control approach using a Y-Thomson spin, is used to detumble the 3U Cubesat with stowed sail and subsequently to 3-axis stabilise the satellite to be ready for the final deployment phase. Minituarised torquer rods, a nano-sized momentum wheel, attitude sensor hardware (magnetometer, sun, earth) developed for this phase will be presented. The final phase will be to deploy and 3-axis stabilise the drag sail normal to the satellite’s velocity vector, using a combined Y-momentum wheel and magnetic controller. The design and performance improvements when using a 2-axis translation stage to adjust the sail centre-of-pressure to satellite centre-of-mass offset, will also be discussed, although for launch risk reasons this stage was not included in the final flight configuration. To accurately determine the drag sail’s attitude during the sunlit part of the orbit, an accurate wide field of view dual sensor to measure both the sun and nadir vector direction was developed for this mission. The calibration results for this new Cubesat sensor ( CubeSense ), will also be presented. [ABSTRACT FROM AUTHOR]

Abstract: Attitude regulation proves to be a challenging problem, when magnetic actuators alone are used as attitude effectors, since they do not provide three independent control torque components at each time instant. In this paper a rigorous proof of global exponential stability is derived for a magnetic control law that leads the satellite to a desired spin condition around a principal axis of inertia, pointing the spin axis toward a prescribed direction in the inertial frame. The technique is demonstrated by means of numerical simulation of a few example maneuvers. An extensive Monte Carlo simulation is performed for random initial conditions, in order to investigate the effect of changes in control law gains. [Copyright &y& Elsevier]