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The escalating proliferation of space debris poses an increasing risk to spinning satellites, elevating the probability of hazardous collisions that can result in severe damage or total loss of functionality. To address this concern, a pioneering inflatable protective structure is employed to ensure the optimal functionality of spinning satellites. Additionally, a multi-body dynamic modeling method based on spring hinge unfolding/spring expansion is proposed to tackle the complex dynamics of spinning satellites with inflatable protective structures during flight. This method enables analysis of the motion parameters of spinning satellites. First, the structural composition of a spinning satellite with inflatable protective structures is introduced and its flight process is analyzed. Then, an articulated spring hinge unfolding model or a spring expansion model using the Newton–Euler method is established to describe the unfolding or expansion of the spinning satellite with inflatable protective structures during flight. Finally, the effects on the motion parameters of a spinning satellite are analyzed through simulation under various working conditions. [ABSTRACT FROM AUTHOR]

A spin stabilized satellite reorientation maneuver in the inertial space is investigated. Spin axis pointing and nutation damping magnetic attitude control algorithms are utilized. Evolutionary equations for the oscillations of a symmetrical satellite near the required position are derived. The exact solution for the wobble amplitude is obtained for the averaged equations of motion. Spin axis attitude angles relative to the required direction are analyzed. Optimal control gain is found for the reorientation maneuver. Theoretical results are verified with numerical simulation. [ABSTRACT FROM AUTHOR]

The escalating proliferation of space debris poses an increasing risk to spinning satellites, elevating the probability of hazardous collisions that can result in severe damage or total loss of functionality. To address this concern, a pioneering inflatable protective structure is employed to ensure the optimal functionality of spinning satellites. Additionally, a multi-body dynamic modeling method based on spring hinge unfolding/spring expansion is proposed to tackle the complex dynamics of spinning satellites with inflatable protective structures during flight. This method enables analysis of the motion parameters of spinning satellites. First, the structural composition of a spinning satellite with inflatable protective structures is introduced and its flight process is analyzed. Then, an articulated spring hinge unfolding model or a spring expansion model using the Newton–Euler method is established to describe the unfolding or expansion of the spinning satellite with inflatable protective structures during flight. Finally, the effects on the motion parameters of a spinning satellite are analyzed through simulation under various working conditions.

Satellite angular motion under the action of the Sdot one-axis magnetic control algorithm is analyzed. Sdot control stabilizes the maximum moment of inertia axis towards the Sun. Evolutionary equations that avoid singularity in the required position are derived. Linearization of equations is performed and new variables that describe the maximum moment of inertia axis oscillations amplitudes are introduced. The resulting equations are suitable for the averaging method application. Evolutionary equations for slow variables are solved. Simplified evolutionary expressions are verified with numerical simulation. [ABSTRACT FROM AUTHOR]

A spin stabilized satellite reorientation maneuver in the inertial space is investigated. Spin axis pointing and nutation damping magnetic attitude control algorithms are utilized. Evolutionary equations for the oscillations of a symmetrical satellite near the required position are derived. The exact solution for the wobble amplitude is obtained for the averaged equations of motion. Spin axis attitude angles relative to the required direction are analyzed. Optimal control gain is found for the reorientation maneuver. Theoretical results are verified with numerical simulation.

Satellite angular motion under the action of the Sdot one-axis magnetic control algorithm is analyzed. Sdot control stabilizes the maximum moment of inertia axis towards the Sun. Evolutionary equations that avoid singularity in the required position are derived. Linearization of equations is performed and new variables that describe the maximum moment of inertia axis oscillations amplitudes are introduced. The resulting equations are suitable for the averaging method application. Evolutionary equations for slow variables are solved. Simplified evolutionary expressions are verified with numerical simulation.

Fast rotating satellite with a rotor is considered. Overall angular momentum of the satellite-rotor system is small. Satellite rotation rate is determined by the rotor angular momentum. This momentum is constant in the nominal attitude mode. Magnetic control system is used to ensure the total angular momentum convergence to zero and spin axis stabilization along the satellite velocity on a circular orbit. Control dipole moment is proposed to provide the necessary equilibrium position. Control torque that resembles the damping law is developed to provide the asymptotic stability. Issues of the control implementation by the magnetic system are considered. • Fast rotating satellite with a rotor is considered. • Spin axis is directed along the satellite orbital velocity. • Satellite-rotor system angular momentum is small. • Magnetorquers provide the rotation and attitude maintenance. [ABSTRACT FROM AUTHOR]

Summary: In this paper, we present the in‐orbit verification of a dynamic link budget approach for the communication link of the Eu:CROPIS (Euglena and Combined Regenerative Organic‐Food Production in Space) satellite. Because of the high satellite spin rates of up to 30 rpm, the satellite antennas radio frequency beam becomes strongly dynamic, consisting of fast amplitude variations and phase rotations. Both effects degrade the link quality and might lead to an abort of the communication. The paper presents an accurate method of how to simulate these dynamic effects and examines the simulation outputs with real in‐orbit measurements. Both results are compared with each other, and deviations are discussed. [ABSTRACT FROM AUTHOR]

Satellites are playing vital role in the astronomical study by scientific community. Whereas making a bigger satellite, catering complex mission requirement for scientific study is not viable due to huge realization duration and launch cost. Configurations of astronomical study satellites are quite different compared to earth observation satellites, in such a way that payloads of former may demand rotating mechanisms. A cost effective spinning satellite configuration is studied to meet the payload rotation requirement for space astronomy exploration. On the other hand, accomplishing the essential requirements is a challenge viz. power generation, stringent payload pointing, attitude updates, telecommand establishment, telemetry links, visibility to data reception ground station for a spinning satellite. This paper is intended to bring out the complexity involved and the systems engineering concepts adopted to achieve the mission goal, mainframe elements redundancy, positive energy balance and high systems reliability to build a cost effective spinning small satellite. [ABSTRACT FROM AUTHOR]

This article presents the design and simulation of the visible band imaging system for small sun synchronous LEO spinning satellite. Unlike conventional remote sensing satellites, which are designed based on three-axis attitude control systems, in this paper, the design is based on a spin stabilized satellite. In a spin-stabilized satellite, the satellite body is spun at a rate about an axis perpendicular to the orbital plane. Satellite imaging payload is based on linear CCD sensor that is placed in parallel with the axis of the satellite rotation. Hardware design is provided as a logic resulting in simplicity as well as more reliability of payload due to no need for programming. Moreover, because of the ICs’ small surface, this payload is resistant against cosmic radiations and in the other words it is rad hard. It also results in a much lower cost comparing to the processor based structures which benefiting rad hard capability. The required parameters to design an imaging payload are calculated which, include integration time, optic aperture, focal length, field of view, image plane dimension, quantitative inefficiencies of detector, optic filter requirements and sampling parameters. For optic design, target radiation, which depends on sun angle, atmosphere passing and target characteristics, is calculated. After this, the CCD readout system, data handling system and the required circuits of these systems are presented in detail. [ABSTRACT FROM AUTHOR]

This paper investigates the use of model-based predictive control for the capture of a multi-degrees-of-freedom object that moves in a somewhat predictable manner, using a deployable manipulator. The study is conducted through both computer simulations and ground-based experiments with the intended application focused on automating the robotic capture of a free-floating and spinning satellite. The present investigation uses an innovative manipulator known as the multi module deployable manipulator system (MDMS), which has been designed and built in our laboratory. The MDMS is used to capture the simulated target satellites. The objective is to evaluate the performance of the predictive controller for the capturing task and to investigate the need and feasibility of incorporating constraints into the controller. Offline solutions to multi parametric quadratic programming (mp-QP) problems are used to create a lookup table so that constrained optimal control decisions can be made in real time. Moreover, a suboptimal formulation of the mp-QP problem has been used to reduce the size of the online lookup table and to mitigate any numerical problems. The results show that when the satellite motion is predicted, the tracking performance is improved. Furthermore, when the physical constraints of the system are formulated into the optimization, the controller becomes aware of its own limitations and the approach toward the satellite is improved by eliminating the possible overshooting of the target [ABSTRACT FROM PUBLISHER]

This work is an investigation on the roots of chaotic attitudinal motion in a class of asymmetrical gyrostat satellites. The result shows that for a class of Kelvin type gyrostat satellite, there is an equivalent rigid spinning satellite with the same attitude dynamics. Finding some constants of motion and eliminating the cyclic coordinates, the rotational kinetic energy is changed to a quadratic form and using Jordan canonical form of the associated inertia tensor and transforming the coordinate system, the Hamiltonian has been changed to those of a rigid satellite. The Hamiltonian has been split into integrable and non-integrable parts. Using Deprit canonical transformation and Andoyer variables the integrable part has been reduced to a one-dimensional form. The reduced Hamiltonian shows that the regular dynamics of the satellite can be chaotic, under the influence of gravitational effects. To demonstrate various attitudinal dynamics of the satellite, a second-order Poincare´ map is employed. This research shows firstly, that the attitudinal dynamics of Kelvin type gyrostat satellites and rigid satellites follow the same dynamical patterns, secondly, for non-linear analysis of dynamics of gyrostat satellite based on the perturbation methods, there is a preferable form for Hamiltonian of the system in the near-integrable fashion and thirdly the chaotic motion is originated from the gravitational field effects that can be suppressed by increasing the attitudinal energy of the satellite in comparison with the translational energy. [Copyright &y& Elsevier]

Near integrability is investigated in a class of gyrostat satellites. The model used consists of a body and a constant speed rotor. The torque applied to the rotor is generated by a potential field. It is shown that there is an equivalent rigid spinning satellite with the same attitude dynamics. The Hamiltonian of the system is converted and reduced into a near integrable form which is suitable for perturbation analysis. [ABSTRACT FROM AUTHOR]

Aryabhata was a spin-stabilised satellite. Attitude determination was required for the evaluation of the stabilisation system and for scientific experiments. Temperature sensors were required for the performance evaluation of the onboard systems and of the thermal control system. A detailed description of the sensors used for these functions in Aryabhata is given. [ABSTRACT FROM AUTHOR]