Fast Attitude Maneuver of a Flexible Spacecraft with Passive Vibration Control Using Shunted Piezoelectric Transducers

Abstract

This paper is concerned with designing a bang-bang control input to perform a quick rotational maneuver of a rigid spacecraft hub connected with flexible appendages. The control design is based on only the rigid body mode making it very simple to design and at the same time achieve the quickest maneuver possible. The induced vibrations are suppressed using piezoelectric transducers bonded to the appendages and connected to an electric circuit with the objective of converting the vibrational energy to electrical energy and then dissipating it using passive electric elements, such as a resistance and an inductor. The proposed control design method is applied to a spacecraft containing a rigid hub and flexible appendages. The attitude control torque is produced using either the reaction wheels contained inside the rigid hub or jet thrusters mounted outside it. The control design process starts with deriving the nonlinear partial differential equations of motion for the spacecraft using Hamilton’s principle which accounts for the electromechanical coupling and the presence of resistive or resistive-inductive circuits. To simplify the analysis, the nonlinear ordinary differential equations of motion are then obtained using the assumed mode method. The effectiveness of the control design method is numerically tested on a spacecraft that is required to perform a quick attitude maneuver and, simultaneously, suppress the induced vibrations. The simulations show a quick and accurate maneuver has been achieved combined with very low levels of vibrations resulting from the reduced coupling between flexible and rigid motions as well as the damping added as a result of the passive shunt circuit. Furthermore, the resistance-inductance shunt circuit is shown to be more effective in damping the vibrations than the resistance shunt circuit.

Finite‐time attitude‐orbit integrated tracking control of flexible spacecraft under input saturation

“…Remark 1. As we can see, orbit dynamics and attitude dynamics are described separately in Equation (12), which means that two different controllers should be designed for orbit and attitude, respectively. What is more, the terms B r T ω and B t T ̇v in the vibration equation bring huge challenges to controller design.…”

“…Next, through the transformation of Equation ( 12), an integrated dynamic model is obtained. For Equation (12), by substituting the vibration equation into orbit and attitude dynamic equations, there are…”

“…For spacecraft with flexible appendages, orbit motion and attitude motion arise the vibration of flexible appendages; vibration in turn affects orbit and attitude stability. Although flexible appendages bring strong coupling and nonlinearity to the system, a wealth of theoretical results related to flexible spacecraft modeling and control have been produced with the joint efforts of the scholars [10][11][12]. Especially in reference [10], the basic theory of attitude dynamic modeling for three-axis stabilized spacecraft with flexible appendages is systematically expounded.…”

Compound Attitude Maneuver and Collision Avoiding Control for a Novel Noncontact Close-Proximity Formation Satellite Architecture

“…With the development of the space camera technology, highresolution imaging under attitude maneuver has been focused on recently [1][2][3]. However, the imaging results could be declined seriously by the huge flexible solar array and the vibration-generating devices [4][5][6].…”

Envelope-Based Variable-Gain Control Strategy for Vibration Suppression of Solar Array Using Reaction Wheel Actuator

“…The ultralow natural frequency of the large-scale flexible appendage f n is the main reason for slow vibration attenuation. The vibration can result from various on-orbit excitations, e.g., the spacecraft's attitude adjusting, the payload pointing, and the spacecraft docking [3][4][5]. Besides, due to the absence of air resistance and weak friction between mechanisms under low gravity, the inherent modal damping ratio ζ n of the appendage is quite small.…”