Three-axis Attitude Control Research Articles

Three-Axis Attitude Control Performance of a Small Satellite Using Control Moment Gyroscope

In this paper, a three-axis attitude control of a small satellite employing control moment gyroscope (CMG) is presented. A four single gimbal control moment gyroscopes (4-SGCMG) cluster has been employed as the actuator and a proportional-derivative (PD) based attitude control approach has been utilized. The singularity robust (SR) steering law has been employed to investigate its capability in avoiding the singularity state of the CMG system with respect to the mission operation by setting the initial gimbal angle further from the singularity state. The presented control law was simulated in the MATLAB® SIMULINK® software. Results obtained from the simulation show the effectiveness of the control algorithm.

Three-axis L1 adaptive attitude control of spacecraft using solar radiation pressure

Based on the [Formula: see text] adaptive control theory, a novel three-axis attitude control system for an axisymmetric spacecraft with uncertain dynamics moving in elliptic orbits, using solar radiation pressure, is derived. The nonaffine-in-control nonlinear spacecraft model includes the gravity gradient torque, the control torque produced by four solar flaps, and external time-varying disturbance moments. For the three-axis attitude control, an [Formula: see text] adaptive control system is designed, which includes a state predictor. A smooth projection algorithm is used to confine the estimated parameters within a desirable set. In the closed-loop system, the designed adaptive law accomplishes three-dimensional attitude control. A special feature of the attitude control system is that it is possible to select large adaptation gains for fast adaptation and to obtain quantifiable performance bounds. Simulation results show that in the closed-loop system, precise roll, yaw, and pitch angle control is accomplished, despite unmodeled nonlinearities, parameter uncertainty, and external disturbance inputs in the model.

ADDASAT Development Plan for the Attitude Control System

As a part of the ORVEAL project, a real scale laboratory model of the inertial platform for ADDASAT is developed by the joint ADDA and UPB teams, ensuring the capability of a three-axis attitude control demonstration in the ground laboratory that simulates weightlessness by low friction bearings. The study is part of the larger ADDA-UPB program for developing the low cost NERVA orbital system for applications in enhanced environmental policies and land resources surveillance. The ORVEAL research is granted by Romanian UEFISCDI financing authority.

BRITE-Constellation: Nanosatellites for precision photometry of bright stars

AbstractBRITE-Constellation (where BRITE stands for BRIght Target Explorer) is an international nanosatellite mission to monitor photometrically, in two colours, brightness and temperature variations of stars brighter than V ≈ 4, with precision and time coverage not possible from the ground. The current mission design consists of three pairs of 7 kg nanosats (hence “Constellation”) from Austria, Canada and Poland carrying optical telescopes (3 cm aperture) and CCDs. One instrument in each pair is equipped with a blue filter; the other, a red filter. The first two nanosats (funded by Austria) are UniBRITE, designed and built by UTIAS-SFL (University of Toronto Institute for Aerospace Studies-Space Flight Laboratory) and its twin, BRITE-Austria, built by the Technical University Graz (TUG) with support of UTIAS-SFL. They were launched on 25 February 2013 by the Indian Space Agency, under contract to the Canadian Space Agency.Each BRITE instrument has a wide field of view (≈ 24 degrees), so up to 15 bright stars can be observed simultaneously in 32 × 32 sub-rasters. Photometry (with reduced precision but thorough time sampling) of additional fainter targets will be possible through on-board data processing. A critical technical element of the BRITE mission is the three-axis attitude control system to stabilize a nanosat with very low inertia. The pointing stability is better than 1.5 arcminutes rms, a significant advance by UTIAS-SFL over any previous nanosatellite.BRITE-Constellation will primarily measure p- and g-mode pulsations to probe the interiors and ages of stars through asteroseismology. The BRITE sample of many of the brightest stars in the night sky is dominated by the most intrinsically luminous stars: massive stars seen at all evolutionary stages, and evolved medium-mass stars at the very end of their nuclear burning phases (cool giants and AGB stars). The Hertzsprung-Russell diagram for stars brighter than mag V=4 from which the BRITE-Constellation sample will be selected is shown in Fig. 1. This sample falls into two principal classes of stars:(1) Hot luminous H-burning stars (O to F stars). Analyses of OB star variability have the potential to help solve two outstanding problems: the sizes of convective (mixed) cores in massive stars and the influence of rapid rotation on their structure and evolution.(2) Cool luminous stars (AGB stars, cool giants and cool supergiants). Measurements of the time scales involved in surface granulation and differential rotation will constrain turbulent convection models.Mass loss from these stars (especially the massive supernova progenitors) is a major contributor to the evolution of the interstellar medium, so in a sense, this sample dominates cosmic “ecology” in terms of future generations of star formation. The massive stars are believed to share many characteristics of the lower mass range of the first generation of stars ever formed (although the original examples are of course long gone).BRITE observations will also be used to detect some Jupiter- and even Neptune-sized planets around bright host stars via transits, as expected on the basis of statistics from the Kepler exoplanet mission. Detecting planets around such very bright stars will greatly facilitate their subsequent characterization. BRITE will also use surface spots to investigate stellar rotation.The following Table summarizes launch and orbit parameters of BRITE-Constellation components.The full version of this paper describing in more detail BRITE-Constellation will be published separately in a journal. The symposium presentation is available at http://iaus301.astro.uni.wroc.pl/program.php

Designing, Manufacturing, and Testing a Self-Contained and Autonomous Nanospacecraft Attitude Control System

AbstractThis paper describes the development of a three-axis attitude control system for nanosatellites and CubeSats, based on commercial electronics and capable to work autonomously for attitude stabilization or controlled from Earth to perform real-time orientation maneuvers. The system contains three microreaction wheels and related drivers, three magnetorquers with their control circuits, three magnetometers, and control electronics. The system is completely independent from the host spacecraft, with all systems included in its structure. For autonomous attitude control, an external sensor is required, providing the orientation of a reference source, such as the sun, eventually evaluated based on solar array outputs, Earth, or stars. The communication interface is based on a serial link and can be used to transfer commands, parameters, and housekeeping data. Miniaturization of the system components and of the whole assembly are among the main innovations introduced with respect to traditional attitude...

Rao-Blackwellized particle filter with vector observations for satellite three-axis attitude estimation and control in a simulated testbed

A Rao-Blackwellized particle filter has been designed and its performance investigated in a simulated three-axis satellite testbed used for evaluating on-board attitude estimation and control algorithms. Vector measurements have been used to estimate attitude and angular rate and, additionally, a pseudo-measurement based on a low-pass filtered time-derivative of the vector measurements has been proposed to improve the filter performance. Conventional extended and unscented Kalman filters, and standard particle filtering have been compared with the proposed approach to gauge its performance regarding attitude and angular rate estimation accuracy, computational workload, convergence rate under uncertain initial conditions, and sensitivity to disturbances. Though a myriad of filters have been proposed in the past to tackle the problem of spacecraft attitude and angular rate estimation with vector observations, to the best knowledge of the authors the present Rao-Blackwellized particle filter is a novel approach that significantly reduces the computational load, provides an attractive convergence rate, and successfully preserves the performance of the standard particle filter when subjected to disturbances.

Small satellite REIMEI for auroral observations

This paper describes the outline and the five years' on-orbit results of the small scientific satellite REIMEI for aurora observations and demonstrations of advanced small satellite technologies. REIMEI is a small satellite with 72 kg mass, and is provided with three-axis attitude control capabilities for aurora observations. REIMEI was launched into a nearly sun synchronous polar orbit on Aug. 23rd, 2005, from Baikonur, Kazakhstan, by Dnepr rocket. REIMEI satellite has been satisfactorily working on the orbit for five years at present as of January, 2011. Three-axis control is achieved with accuracy of 0.1° (3 σ). Multi-spectrum images of aurora are taken with 8 Hz rate and 2 km spatial resolution to investigate the aurora physics. REIMEI is performing the simultaneous observation of aurora images and particle measurements. REIMEI indicates that even a small satellite launched as a piggy-back can successfully perform unique scientific mission purposes.

Design, Development and Operation of a Laboratory Pulsed Plasma Thruster for the First Time in West Asia

Although the pulsed plasma thruster (PPT) was first utilized on a space mission in 1964, after more than four decades, it is still a space-rated technology which has performed various propulsion tasks, from station-keeping to three-axis attitude control for a variety of former missions. With respect to the rapid growth in the small satellite community and the growing interest for smaller satellites in recent years, the PPT is one of the promising electric propulsion devices for small satellites (e.g., CubeSats) due to the following advantages: simplicity, lightweight, robustness, low power consumption, low production cost and small dimensions. Therefore, a laboratory benchmark rectangular breech-fed pulsed plasma thruster using a self-inductor as a coupling element was designed, developed and successfully tested in a bell-type vacuum chamber at 10-4 Pa for the first time in west Asia (Iran). The PPT has been tested using a 35 μF, 2.5 kV oil-filled capacitor, producing an impulse bit varying from 300 μN-s to 1.3 mN-s at a maximum specific impulse of 1100 s. As a result a research program in Iran was initiated for working on PPTs and the miniaturization of PPTs while increasing the performance parameters. The present paper briefly reviews the PPT design and development.

Nonlinear filtering in a simulated three-axis satellite attitude estimation and control tested

Nonlinear filtering in a simulated three-axis satellite attitude estimation and control tested

Spacecraft Attitude Control Using a Double-Gimbal Control Moment Gyro

This paper presents a novel attitude control scheme for a spacecraft. It involves the use of a double-gimbal control moment gyro (CMG), wherein the rotational speed of the momentum wheel is kept constant. Tilting the spin axis of the momentum wheel around either an inner or an outer gimbal axis can generate attitude control torques in one rotational direction. In order to compensate for the absence of control torque in the other rotational direction, a combination of rotations of the inner and outer axes is designed to generate a cyclic motion about the spin axis of the momentum wheel resulting in three-axis attitude control. Analytical and numerical algorithms are developed to compute the control inputs, i.e., the tilt angles of the inner and outer gimbal axes, which are responsible for the desired change in attitude. The effectiveness of the proposed control scheme is verified by numerical simulations.

Fuzzy bang-bang relay controller for satellite attitude control system

Two level bang-bang controllers are generally used in conjunction with the thrust reaction actuator for spacecraft/satellite attitude control. These controllers are fast acting and dispense time dependent; full or no thrust-power to control the satellite attitude in minimum time. A minimum time-fuel attitude control system extends the life of a satellite and is the main focus of this paper. Fuzzy controllers are favored for satellite control due to their simplicity and good performance in terms of fuel saving, absorbing non-linearities and uncertainties of the plant. A fuzzy controller requires a soft fuzzy engine, and a hardware relay to accomplish bang-bang control action. The work in this paper describes a new type of fuzzy controller in which the hardware relay action is configured in the soft fuzzy engine. The new configuration provides fuzzy decision-making flexibility at the inputs with relay like two-level bang-bang output. The new fuzzy controller is simulated on a three-axis satellite attitude control platform and compared with conventional a fuzzy controller, sliding mode controller and linear quadratic regulator. The result shows that the proposed controller has minimum-time response compared to other controllers. Inherent chattering associated with a two-level bang-bang controller produces undesirable low amplitude frequency limit cycles. The chattering can be easily stopped in the proposed fuzzy bang-bang relay controller, hence adding multi-functionality to its simple design.

Attitude dynamics and control of satellites orbiting rotating asteroids

The paper focuses on the attitude dynamics and control of satellites orbiting a rotating asteroid. The general formulation of the satellite equations of motion in an equatorial eccentric orbit is obtained through the Lagrangian method. The linearized system model is derived and the stability analysis is presented. The control laws for three-axis attitude control of satellites are developed and a closed-form solution of the system is derived. For an illustration of the linear analysis followed by the numerical simulation of the governing nonlinear equations of motion of the satellite, several asteroids (Eros, Castalia, Vesta, Ida and Gaspra) are considered. Attitude resonances for satellites in retrograde orbits are found. The results of the linear system model compare well with the corresponding system nonlinear equations of motion. The proposed controllers are successful in stabilizing the attitude of the satellites even in the presence of high attitude disturbances and orbital eccentricities.

Dynamic Control Allocation for Shaping Spacecraft Attitude Control Command

For spacecraft attitude control, reaction wheel (RW) steering laws with more than three wheels for three-axis attitude control can be derived by using a control allocation (CA) approach. 1-2 The CA technique deals with a problem of distributing a given control demand to available sets of actuators. 3-4 There are many references for CA with applications to aerospace systems. For spacecraft, the control torque command for three body-fixed reference frames can be constructed by a combination of multiple wheels, usually four-wheel pyramid sets. Multi-wheel configurations can be exploited to satisfy a body-axis control torque requirement while satisfying objectives such as minimum control energy. 1-2 In general, the reaction wheel steering laws determine required torque command for each wheel in the form of matrix pseudo-inverse. In general, the attitude control command is generated in the form of a feedback control. The spacecraft body angular rate measured by gyros is used to estimate angular displacement also.? Combination of the body angular rate and attitude parameters such as quaternion and MRPs(Modified Rodrigues Parameters) is typically used in synthesizing the control command which should be produced by RWs.¹ The attitude sensor signals are usually corrupted by noise; gyros tend to contain errors such as drift and random noise. The attitude determination system can estimate such errors, and provide best true signals for feedback control.? Even if the attitude determination system, for instance, sophisticated algorithm such as the EKF(Extended Kalman Filter) algorithm?, can eliminate the errors efficiently, it is quite probable that the control command still contains noise sources. The noise and/or other high frequency components in the control command would cause the wheel speed to change in an undesirable manner. The closed-loop system, governed by the feedback control law, is also directly affected by the noise due to imperfect sensor characteristics. The noise components in the sensor signal should be mitigated so that the control command is isolated from the noise effect. This can be done by adding a filter to the sensor output or preventing rapid change in the control command. Dynamic control allocation(DCA), recently studied by Harkegard, is to distribute the control command in the sense of dynamics⁴: the allocation is made over a certain time interval, not a fixed time instant. The dynamic behavior of the control command is taken into account in the course of distributing the control command. Not only the control command requirement, but also variation of the control command over a sampling interval is included in the performance criterion to be optimized. The result is a control command in the form of a finite difference equation over the given time interval.⁴ It results in a filter dynamics by taking the previous control command into account for the synthesis of current control command. Stability of the proposed dynamic control allocation (CA) approach was proved to ensure the control command is bounded at the steady-state. In this study, we extended the results presented in Ref. 4 by adding a two-step dynamic CA term in deriving the control allocation law. Also, the strict equality constraint, between the virtual and actual control inputs, is relaxed in order to construct control command with a smooth profile. The proposed DCA technique is applied to a spacecraft attitude control problem. The sensor noise and/or irregular signals, which are existent in most of spacecraft attitude sensors, can be handled effectively by the proposed approach.

Variable structure attitude manoeuvre control and active vibration damping of three-axis-stabilized flexible spacecraft

This paper presents a dual-stage control system design method for the three-axis-rotational manoeuvre and vibration stabilization of a spacecraft with flexible appendages embedded with piezoceramics as the sensor and actuator. In this design approach, the attitude control system and vibration suppression were designed separately using a lower-order model. The design of the attitude controller was based on the variable structure control (VSC) theory, leading to a discontinuous control law. This controller accomplishes asymptotic attitude maneouvring in the closed-loop system and is insensitive to the interaction of elastic modes and uncertainty in the system. The advantage of this dual-stage approach is that once the desired orientation is attained under the VSC law, only the elastic motion is governed by the decoupled linear second-order systems describing elastic dynamics. To this end, modal velocity feedback and optimal positive position feedback control methods are presented and compared for actively damping the elastic oscillations using piezoelectric materials as an additional sensor and actuator bonded on the surface of the flexible appendages. In addition, a special configuration of actuators for three-axis attitude control is also investigated: the pitch attitude controlled by a momentum wheel and the roll/yaw control achieved by on-off thrusters, which is modulated by the pulse width-pulse frequency modulation technique to construct a proper control torque history. Numerical simulations are performed to show that the rotational manoeuvre and vibration suppression are accomplished in spite of the presence of disturbance torque and parameter uncertainty.

Predictive control with mean derivative based neural euler integrator dynamic model

Neural networks can be trained to get internal working models in dynamic systems control schemes. This has usually been done designing the neural network in the form of a discrete model with delayed inputs of the NARMA type (Non-linear Auto Regressive Moving Average). In recent works the use of the neural network inside the structure of ordinary differential equations (ODE) numerical integrators has also been considered to get dynamic systems discrete models. In this paper, an extension of this latter approach, where a feed forward neural network modeling mean derivatives is used in the structure of an Euler integrator, is presented and applied in a Nonlinear Predictive Control (NPC) scheme. The use of the neural network to approximate the mean derivative function, instead of the dynamic system ODE instantaneous derivative function, allows any specified accuracy to be attained in the modeling of dynamic systems with the use of a simple Euler integrator. This makes the predictive control implementation a simpler task, since it is only necessary to deal with the linear structure of a first order integrator in the calculations of control actions. To illustrate the effectiveness of the proposed approach, results of tests in a problem of orbit transfer between Earth and Mars and in a problem of three-axis attitude control of a rigid body satellite are presented.

Three-axis attitude control using combined gravity-gradient and solar pressure

An attitude satellite controller that combines passive gravity-gradient torque with solar pressure torque is proposed for three-axis stabilization of a micro-satellite. A dumbbell satellite configuration is chosen to increase the gravity-gradient torque and a three-axis stabilization is pursued with the help of small solar sails. Solar pressure task is to help the gravity-gradient torque and not to oppose it. The proposed controller is applied to different test cases: first, for satellite stabilization during the tumble motion, due to a nonperfect separation manoeuvre from the launcher last stage; second, the attitude control performances are studied for different eccentric orbits. The results are given in terms of settling time and maximum amplitude oscillations.

Three-axis air-bearing based platform for small satellite attitude determination and control simulation

A frictionless environment simulation platform, utilized for accomplishing three-axis attitude control tests in small satellites, is introduced. It is employed to develop, improve, and carry out objective tests of sensors, actuators, and algorithms in the experimental framework. Different sensors (i.e. sun, earth, magnetometer, and an inertial measurement unit) are utilized to assess three-axis deviations. A set of three inertial wheels is used as primary actuators for attitude control, together with three mutually perpendicular magnetic coils intended for desaturation purposes, and as a backup control system. Accurate balancing, through the platform’s center of mass relocation into the geometrical center of the spherical air-bearing, significatively reduces gravitational torques, generating a virtually torque-free environment. A very practical balancing procedure was developed for equilibrating the table in the local horizontal plane, with a reduced final residual torque. A wireless monitoring system was developed for on-line and post-processing analysis; attitude data are displayed and stored, allowing properly evaluate the sensors, actuators, and algorithms. A specifically designed onboard computer and a set of microcontrollers are used to carry out attitude determination and control tasks in a distributed control scheme. The main components and subsystems of the simulation platform are described in detail.

Robust and optimal attitude stabilization of spacecraft with external disturbances

This paper presents a robust and optimal three-axis attitude control scheme for the rigid body motion with external disturbances, in which a worst-case design that the external disturbance signal maximizes the same performance index that the control input minimizes is assumed. After presenting a robust control law for the three-axis attitude control of the rigid body motion with external disturbances, the optimality property of the robust control law is investigated based on the minimax approach and the inverse optimal approach. A numerical example is given to illustrate the theoretical result presented in this paper.

Three-axis Attitude Control for Flexible Spacecraft by Lyapunov Approach under Gravity Potential

Attitude control law synthesis for the three-axis attitude maneuver of a flexible spacecraft model is presented in this study. The basic idea is motivated by previous works for the extension into a more general case. The new case includes gravitational gradient torque which has significant effect on a wide range of low earth orbit missions. As the first step, the fully nonlinear dynamic equations of motion are derived including gravitational gradient. The control law design based upon the Lyapunov approach is attempted. The Lyapunov function consists of a weighted combination of system kinetic and potential energy. Then, a set of stabilizing control law is derived from the basic Lyapunov stability theory. The new control law is therefore in a general form partially validating the previous work in some sense.

Rotational motion control of a spacecraft

This note describes a systematic procedure for the control synthesis of a rigid spacecraft using the energy shaping method. The geometric concept of a mechanical system in a coordinate-independent form is used to derive a control algorithm for the Euler-Poincare equations. The main result of this note is a specialization of the method on the unit quaternion group. This note is concluded with the examples of the potential functions and implementation for the three-axis attitude control problem.