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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]

Spacecraft attitude control solutions typically are torque-level algorithms that simultaneously control both the attitude and angular velocity tracking errors. In contrast, robotic control solutions are kinematic steering commands where rates are treated as the control variable, and a servo-tracking control subsystem is present to achieve the desired control rates. In this paper kinematic attitude steering controls are developed where an outer control loop establishes a desired angular response history to a tracking error, and an inner control loop tracks the commanded body angular rates. The overall stability relies on the separation principle of the inner and outer control loops which must have sufficiently different response time scales. The benefit is that the outer steering law response can be readily shaped to a desired behavior, such as limiting the approach angular velocity when a large tracking error is corrected. A Modified Rodrigues Parameters implementation is presented that smoothly saturates the speed response. A robust nonlinear body rate servo loop is developed which includes integral feedback. This approach provides a convenient modular framework that makes it simple to interchange outer and inner control loops to readily setup new control implementations. Numerical simulations illustrate the expected performance for an aggressive reorientation maneuver subject to an unknown external torque. [ABSTRACT FROM AUTHOR]

The paper deals with a satellite in a circular near-Earth orbit, perturbed due to J 2 Earth's oblateness. The satellite interacts with the geomagnetic field by the moments of Lorentz and magnetic forces. The octupole approximation of the Earth's magnetic field is used. The possibility of electrodynamical attitude control for the satellite's stabilization in the orbital frame is analyzed. Once the problem of electrodynamical compensation of disturbing torque is solved, we can obtain the control algorithms for the satellite electromagnetic parameters which allows to stabilize the satellite attitude position in the orbital frame in the presence of disturbing gravity gradient torque. The total stability of the satellite programmed motion is proved analytically and verified by computer modeling. [ABSTRACT FROM AUTHOR]

Abstract: Feasibility of achieving three axis attitude stabilization using a single thruster is explored in this paper. Torques are generated using a thruster orientation mechanism with which the thrust vector can be tilted on a two axis gimbal. A robust nonlinear control scheme is developed based on the nonlinear kinematic and dynamic equations of motion of a rigid body spacecraft in the presence of gravity gradient torque and external disturbances. The spacecraft, controlled using the proposed concept, constitutes an underactuated system (a system with fewer independent control inputs than degrees of freedom) with nonlinear dynamics. Moreover, using thruster gimbal angles as control inputs make the system non-affine (control terms appear nonlinearly in the state equation). This necessitates the control algorithms to be developed based on nonlinear control theory since linear control methods are not directly applicable. The stability conditions for the spacecraft attitude motion for robustness against uncertainties and disturbances are derived to establish the regions of asymptotic 3-axis attitude stabilization. Several numerical simulations are presented to demonstrate the efficacy of the proposed controller and validate the theoretical results. The control algorithm is shown to compensate for time-varying external disturbances including solar radiation pressure, aerodynamic forces, and magnetic disturbances; and uncertainties in the spacecraft inertia parameters. The numerical results also establish the robustness of the proposed control scheme to negate disturbances caused by orbit eccentricity. [Copyright &y& Elsevier]

Abstract: The attitude dynamics and manoeuvre survey in this paper is performed for a mission scenario involving a penetrator-type spacecraft: an axisymmetric prolate spacecraft spinning around its minor axis of inertia performing a spin axis reorientation manoeuvre. In contrast to most existing spacecraft only one attitude control thruster is available, providing a control torque perpendicular to the spin axis. Having only one attitude thruster on a spinning spacecraft could be preferred for spacecraft simplicity (lower mass, lower power consumption etc.), or it could be imposed in the context of redundancy/contingency operations. This constraint does yield restrictions on the thruster timings, depending on the ratio of minor to major moments of inertia among other parameters. The Japanese Lunar-A penetrator spacecraft proposal is a good example of such a single-thruster spin-stabilised prolate spacecraft. The attitude dynamics of a spinning rigid body are first investigated analytically, then expanded for the specific case of a prolate and axisymmetric rigid body and finally a cursory exploration of non-rigid body dynamics is made. Next two well-known techniques for manoeuvring a spin-stabilised spacecraft, the Half-cone/Multiple Half-cone and the Rhumb line slew, are compared with two new techniques, the Sector-Arc Slew developed by Astrium Satellites and the Dual-cone developed at Surrey Space Centre. Each technique is introduced and characterised by means of simulation results and illustrations based on the penetrator mission scenario and a brief robustness analysis is performed against errors in moments of inertia and spin rate. Also, the relative benefits of each slew algorithm are discussed in terms of slew accuracy, energy (propellant) efficiency and time efficiency. For example, a sequence of half-cone manoeuvres (a Multi-half-cone manoeuvre) tends to be more energy-efficient than one half-cone for the same final slew angle, but more time-consuming. As another example, the new Sector-Arc Slew and Dual-cone techniques are designed to overcome a specific restriction on attainable slew angle that is associated with the half-cone manoeuvre, giving one additional degree of freedom for designers to fine-tune. [Copyright &y& Elsevier]