Reaction Wheel Torque Research Articles

Coupled translation and attitude tracking control for multi-satellite electromagnetic formation flight based on dual quaternion

Electromagnetic formation flight technology utilizes electromagnetic force and torque between satellites to control satellites’ relative position and attitude without consuming propellant, thus having broad application prospects. However, the dynamics of electromagnetic formation flying are nonlinear and strongly coupled, posing challenges to the 6-DOF control of electromagnetic formation and the magnetic dipole allocation of satellites. The frequency division multiplexing method can approximate decoupling formation dynamics by loading multi-frequency AC carriers onto satellite magnetic dipoles and is likely to achieve more control functions. A design method for AC carriers based on frequency division multiplexing is proposed to decouple formation dynamics while controlling electromagnetic force and torque. This paper further provides an analytical solution for the magnetic dipoles of the electromagnetic force/torque equations. In addition, in response to external disturbances, model uncertainty, and the inability to obtain relative velocity and angular velocity information of the formation system, this paper establishes a 6-DOF coupling model based on dual quaternion and implements tracking control using the active disturbance rejection control method. Based on the internal force characteristics of the electromagnetic formation, an expression for the equilibrium distribution of reaction wheel torque is derived to avoid some reaction wheels’ rapid accumulation of angular momentum. Numerical simulation results show the effectiveness of the proposed control method.

Reaction Wheel Torque Distribution Algorithm for Enhancing the Agility of the Spacecraft using Null motion

Earth observation satellites require rapid attitude maneuvering and stringent pointing accuracy for imaging payloads. Four numbers of Reaction Wheel (RW) are usually configured in standard tetrahedral configuration for attitude manoeuvring as well as for fine pointing. The Reaction wheel maximum torque is one of the critical parameters that determine the agility of the spacecraft. Three axis attitude control torque signal is distributed to 4 Reaction wheels using standard pseudo inverse solution which minimizes the sum of the square of the individual wheel torque rather than the maximum of the individual wheel torque. The pseudo inverse solution leads to wheel saturation even when the total momentum does not reach the envelope and there by limiting the agility of the spacecraft. In this paper, new approach is adopted to optimally utilize the full capacities of RWs and thereby enhancing the agility of the spacecraft. The new algorithm is based on the null motion of the reaction wheel torque alignment matrix. The proper choice of the null vector results in equal wheel torque demand when combining with the pseudo inverse solution. The maneuver & control algorithms are integrated together with new distribution method and have been demonstrated for typical rest to rest and rate to rate maneuver.

PocketQube Pico-Satellite Attitude Control: Implementation and Testing

Attitude control for CubeSats and small satellites has been widely researched. Nonetheless, most of the literature is based on simulations, with limited availability of actual implementation results. This article tries to fill this gap by highlighting implementation difficulties and proposing appropriate workable solutions. The difficulty of direct reaction wheel (RW) torque feedback for attitude control of a 1p PocketQube (PQ) pico-satellite is noted and addressed. A simple control strategy implementable on a microcontroller is then proposed, showing how direct RW torque control can be achieved through angular velocity measurement without using derivative feedback. A linear quadratic Gaussian controller with integral action in quaternion formulation is used for attitude control, and a reference input boundary caused by the use of the standard integral action scheme is highlighted. A simple solution is then proposed, showing how using the error quaternion removes this boundary. An extended Kalman filter in quaternion formulation is used for attitude determination by fusing data from three sensors, namely, a magnetometer, solar cells, and a gyroscope. An accelerometer is used in place of the solar cells for simplified laboratory testing. Magnetometer error due to surrounding interference is addressed through calibration. A string-suspended model of the UoMBSat-1 PocketQube pico-satellite is used for attitude control testing. Preliminary single-axis tests are performed using a PID controller to determine the satellite parameters and actuators verification. The proposed linear quadratic Gaussian controller is then tested and validated experimentally in the laboratory setting.

Maximizing agility envelopes for reaction wheel spacecraft

Spacecraft agility is limited by the maximum torque that reaction wheels can provide. Therefore, a reaction wheel array is typically configured to maximize the inscribed sphere of the reaction wheel torque envelope. Agility is then determined by dividing the spherical torque by the maximum principal inertia. This industry standard approach can severely underestimate the true capability of an attitude control system. An agility envelope considers the reaction wheel torque envelope along with the spacecraft inertia tensor. The agility envelope can therefore be used as a means to quantify the conservatism associated with the standard approach in order to improve slew performance of a conventional attitude control system without the need for larger, more costly hardware or new control algorithms. This paper, presents a simple approach for constructing the agility envelope of a reaction wheel attitude control system. The agility envelope is applied to determine design curves for limits on angular acceleration and rate for maneuver design and for finding the reaction wheel skew angles that maximize agility for a given spacecraft configuration. A surprising result is the observation that maximizing the inscribed sphere of the reaction wheel torque envelope does not, in general, optimize agility.

GOES-R active vibration damping controller design, implementation, and on-orbit performance

GOES-R series spacecraft feature a number of flexible appendages with modal frequencies below 3.0 Hz which, if excited by spacecraft disturbances, can be sources of undesirable jitter perturbing spacecraft pointing. To meet GOES-R pointing stability requirements, the spacecraft flight software implements an Active Vibration Damping (AVD) rate control law which acts in parallel with the nadir point attitude control law. The AVD controller commands spacecraft reaction wheel actuators based upon Inertial Measurement Unit (IMU) inputs to provide additional damping for spacecraft structural modes below 3.0 Hz which vary with solar wing angle. A GOES-R spacecraft dynamics and attitude control system identified model is constructed from pseudo-random reaction wheel torque commands and IMU angular rate response measurements occurring over a single orbit during spacecraft post-deployment activities. The identified Fourier model is computed on the ground, uplinked to the spacecraft flight computer, and the AVD controller filter coefficients are periodically computed on-board from the Fourier model. Consequently, the AVD controller formulation is based not upon pre-launch simulation model estimates but upon on-orbit nadir point attitude control and time-varying spacecraft dynamics. GOES-R high-fidelity time domain simulation results herein demonstrate the accuracy of the AVD identified Fourier model relative to the pre-launch spacecraft dynamics and control truth model. The AVD controller on-board the GOES-16 spacecraft achieves more than a ten-fold increase in structural mode damping for the fundamental solar wing mode while maintaining controller stability margins and ensuring that the nadir point attitude control bandwidth does not fall below 0.02 Hz. On-orbit GOES-16 spacecraft appendage modal frequencies and damping ratios are quantified based upon the AVD system identification, and the increase in modal damping provided by the AVD controller for each structural mode is presented. The GOES-16 spacecraft AVD controller frequency domain stability margins and nadir point attitude control bandwidth are presented along with on-orbit time domain disturbance response performance.

Calibration of reaction wheel torque measuring system based on wavelet methods

Calibration of reaction wheel torque measuring system based on wavelet methods

H∞ control of a gravity gradient stabilised satellite

This paper describes the attitude control of a small gravity gradient and reaction wheel stabilised satellite. Gravity gradient stabilised satellite have limited stability and pointing capabilities. Thus, H ∞ control has been applied, in order to improve the accuracy of the attitude control. The synthesis of a suitable H ∞ controller, using a composite weighting function is described. The trade-off between reaction wheel torque and possible disturbance rejection capability using H ∞ control is discussed. Also, the promising results using H ∞ control with antireset windup compensation are presented.

Singularity Avoidance Using Null Motion and Variable-Speed Control Moment Gyros

A variable speed control moment gyroscope (VSCMG) null motion steering law is introduced that continuously attempts to minimize the condition number of the control in uence matrix. By doing so, the gimbal angles are rearranged to less singular conŽ gurations without exerting a torque onto the spacecraft. By allowing the reaction wheel speeds to be variable in this steering law, more general reconŽ gurations are possible than what are possible with conventionalCMG. No a priori calculations of preferred sets of gimbal angles are necessary with this method. Numerical studies show that superimposing this VSCMG null motion on the VSCMG steering law can result in a drastic reduction in the required reaction wheel power consumption when operating near CMG singular gimbal conŽ gurations.The reaction wheel torque required by this VSCMG steering law is typically very small and achievable with existing CMG hardware.

Space Telescope Antenna Pointing System Analysis and Test

Space Telescope, ST, carrying five astronomy instruments, is a long life spacecraft designed to achieve and maintain ultra fine stabilization jitter performance. The RMS jitter specification for pointing performance is 0.007 arc-sec, 1 sigma, during periods of scientific observation. Spacecraft attitude control it achieved by on-board computer processing of attitude and rate sensor data to generate reaction wheel torque commands. The Space Telescope Antenna Pointing System, APS, is used to position a High Gain Antenna, MGA, for relaying science and engineering data from the spacecraft to the ground station.