Small Satellite Attitude Control Research Articles

Research on rapid batch test technology of small satellite attitude and orbit control system

The attitude and orbit control sub-system of small satellites is an important sub-system to ensure the stable operation of satellites in orbit, and its testing in the satellite factory development stage runs through the whole testing process. Most of the traditional attitude and orbit control sub-system tests require the cooperation of many people to complete the performance tests of various components and sensitive devices, which also leads to the complicated and time-consuming testing process of the sub-system. At present, the mass production of small satellites is developing rapidly. In order to meet the urgent needs of small satellite mass production, the traditional attitude control subsystem test equipment and methods can’t meet the requirements of satellite mass production mode. In order to meet the needs of small satellite mass development, this paper studies the rapid batch test technology of small satellite attitude and orbit control systems. A fast pipeline test system which can support the performance and dynamics simulation of various components of a multi-satellite attitude and orbit control subsystem is studied. Based on the architecture of real-time communication between the upper computer and lower computer, the system can realize the integration and rapid testing of component performance and dynamic simulation. A continuous dynamic simulation test of 2140 seconds was carried out for the cruise mode and attitude maneuver mode of a certain type of satellite in orbit. During the test, various components and sensors of the attitude and orbit control subsystem operate correctly and can support the stable operation of the required attitude in various modes of the satellite, which can verify the correctness of the interface and performance of the satellite attitude and orbit control subsystem. The research results greatly reduce the input of manpower and material resources, and greatly improve the test efficiency. It lays a foundation for the rapid batch test of the small satellite attitude control subsystem.

On-orbit performance analysis of spinning spacecraft magnetic control laws. Application to the UPMSat-2 mission

This paper studies the performance of a new magnetic attitude control law developed for the UPMSat-2 satellite. The UPMSat-2 control law, which is a modification of the B-dot, allows the satellite to align one axis with the normal to the orbital plane and achieve a target angular velocity, without requiring attitude determination. The study evaluates the control law’s effectiveness using the first housekeeping data received since the satellite’s launch in September 2020. To verify the control’s on-orbit performance, data from sun sensors and magnetometers, along with thermal analysis of the satellite external temperatures, are used to confirm that the satellite remains at its design attitude and angular velocity. The study concludes that the proposed control law is a simple and effective alternative for small satellite attitude control.

Intelligent robust adaptive attitude controller for small satellites

This work addresses the issue of small-satellite attitude control. This satellite’s Attitude-Control-System (ACS) is managed by an intelligent controller built on Adaptive Neuro-Fuzzy Inference System (ANFIS). ANFIS control approach is designed to improve the satellite’s angular mobility along its axes. Two simulation scenarios have been tackled. In the first scenario, ACS system efficiency was evaluated and contrasted with ACS based on a PD controller optimized by a genetic-algorithm for the nonlinear system model and a Linear-Quadratic-Regulator (LQR) in the existence of input environmental disturbances. The simulation outcomes demonstrate that the proposed control technique successfully addresses the mobility issue of a small satellite. Inaccuracies in the satellite’s parameters, torque noise brought on by the vibration of the actuators, and external disturbances were added in the second scenario. The proposed ANFIS controller has demonstrated robust, reliable performance for control of the ACS.

Logic-based gimbal compensator for the double-gimbal scissored-pair control moment gyroscope using magnetic torquers

Control Moment Gyroscopes (CMGs) are effective torque-producing actuators for satellite missions requiring agile maneuvering and accurate pointing. CMGs can develop relatively large torques compared to other momentum storage actuators like reaction wheels; however, using CMGs for three-axis attitude control is complicated by inherent mechanical singularities associated with CMGs. The singularity effect is worsened by common disturbances, such as aerodynamic or magnetic torques, which influence the CMG momentum state. A recent CMG design known as the Double-Gimbal Scissor-Pair CMG (DGSPCMG) has shown promise for use in small satellites because it has manageable singular conditions and has the potential to achieve three axis control using only two flywheels. This paper proposes a novel extension of the DGSPCMG control law and supplements it with a classical sliding-mode attitude controller to allow for the DGSPCMG momentum to be managed such that the gimbals may be restored to desirable positions by way of magnetic torques. The proposed solution achieves agile maneuvering by differentiating when singularity escape is possible through a steering control law and when magnetic torques are necessary for gimbal compensation in the DGSPCMG. The effectiveness of the control scheme is demonstrated by numerical simulations in MATLAB which subject the design to a detailed space environment. For the simulation conditions used in this research, the results indicate that the proposed control scheme would be an effective solution for attitude control of small satellites using a DGSPCMG in Low-Earth Orbit.

Small-Satellite Attitude Control Using Continuous Sinusoids With Strict Amplitude Constraints

This article considers attitude control of a rigid body (e.g., small satellite) with internal rotating-mass actuators that, unlike reaction wheels, cannot perform unrestricted rotations. Instead, the rotational stroke of each actuator is limited to <inline-formula xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"> <tex-math notation="LaTeX">$[-\alpha,\alpha]$ </tex-math></inline-formula> rad, where <inline-formula xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"> <tex-math notation="LaTeX">$\alpha &gt; 0$ </tex-math></inline-formula> . We present two attitude-feedback control methods. The first method addresses setpoint tracking of constant attitude commands, whereas the second approach addresses command following of time-varying attitude commands. Both methods use control signals that are continuous and piecewise sinusoidal but whose derivatives can contain discontinuities. The main analytic results show that these attitude-feedback controls achieve asymptotic setpoint tracking for a constant attitude command and approximate command following for a time-varying attitude command. The results also show that the controls satisfy the strict actuator amplitude constraint. Each control method is demonstrated in numerical simulations of a small satellite in deep space. Finally, we present single-axis closed-loop attitude control experiments for a small-satellite system on an air bearing.

An AEFA-Based Optimum Design of Fuzzy PID Controller for Attitude Control Flywheel with BLDC Motor

A new method for optimizing the fuzzy PID controller, based on an artificial electric field algorithm (AEFA), is proposed in this paper, aiming at improving the stability indicator of the Brushless DC (BLDC) motor for the small satellite attitude control flywheel. The BLDC motor is the basic part of the small satellite attitude control flywheel. In order to accurately control the attitude of the small satellite, a good motor control system is very important. Firstly, the mathematical model of the BLDC motor is established and the BLDC motor speed control system using traditional PID control is designed. Secondly, considering that the small satellite speed control system is a nonlinear system, a fuzzy PID control is designed to solve the shortcomings of the fixed parameters of the traditional PID control. Finally, we find that the control accuracy of the fuzzy PID control will change with the range of the input. Therefore, we introduce the AEFA to optimize fuzzy PID to achieve high-precision attitude control of small satellites. By simulating the BLDC motor system, the proposed fuzzy PID controller based on AEFA is compared with the traditional PID controller and the fuzzy PID controller. Results from different controllers show that the proposed control method could effectively reduce steady state error. In addition, the proposed fuzzy PID–AEFA controller has the better anti-jamming capability.

Minimum-Time Attitude Maneuver and Robust Attitude Control of Small Satellite Mounted with Data Relay Communication Antenna

This paper proposes a nonlinear control method for carrying out Minimum-time satellite attitude maneuver and antenna motion which have robustness against model uncertainty. In recent years, small Earth observation satellites have been utilized and expected to maneuver rapidly in missions such as multi-target acquisition. On the other hand, small satellites need to send the observation data to ground station. Recently, small Earth observation satellites acquire high-resolution data, resulting in an increase in the time required for data communication. Thus, small satellites need to use inter-orbit communication link through Data Relay test satellite sending data from Data Relay communication (DRC) antenna. In conventional operations, the antenna motion is implemented after satellite attitude maneuver. However, this method has a time delay between the completion of the attitude maneuver and the start of data communication. The purpose of this study is to extend the time of earth observation and data communication by carrying out satellite maneuver and antenna motion concurrently. Because small satellite mounted with DRC antenna has large mass ratio of the antenna, we cannot ignore time variability of the moment of inertia of the whole system and reaction torque generated by antenna motion. Hence, in order to take the influence of the antenna motion into consideration, we combine a satellite attitude control system and an antenna drive system into one control system by governing equations and constructing the optimal control problem. We convert the optimal control problem into a NLP by discretizing the control input (a series of pulses) to minimize the final time of the total maneuver that includes the antenna adjustment. In addition, it is considered that a model uncertainty and unknown disturbance occurs in real space. Thus, we have to design feedback controller to secure robustness in model error and unknown disturbance. Accordingly in order to propose a nonlinear control method for carrying out minimum-time satellite attitude maneuver and antenna motion which have robustness against model uncertainty and unknown disturbance, we calculate a reference attitude by application of optimal control input torque to ideal satellite model and design servo controller by using state-dependent Riccati equation (SDRE) control method in order to track time-variant reference attitude.

On verifying magnetic dipole moment of a magnetic torquer by experiments

Magnetic torquers are used for the attitude control of small satellites, such as CubeSats with Low Earth Orbit (LEO). During the design of magnetic torquers, it is necessary to confirm if its magnetic dipole moment is enough to control the satellite attitude. The magnetic dipole moment can affect the detumbling time and the satellite rotation time. In addition, it is also necessary to understand how to design the magnetic torquer for operation in a CubeSat under the space environment at LEO. This paper reports an investigation of the magnetic dipole moment and the magnetic field generated by a circular air-coil magnetic torquer using experimental measurements. The experiment testbed was built on an air-bearing under a magnetic field generated by a Helmholtz coil. This paper also describes the procedure to determine and verify the magnetic dipole moment value of the designed circular air-core magnetic torquer. The experimental results are compared with the design calculations. According to the comparison results, the designed magnetic torquer reaches the required magnetic dipole moment. This designed magnetic torquer will be applied to the attitude control systems of a 1U CubeSat satellite in the project “KNACKSAT.”

Electrospray Thrusters as Precise Attitude Control Actuators for Small Satellites

The purpose of this work is to describe the use of electrospray propulsion devices as miniaturized actuators for precise attitude control of small satellites, including CubeSats. Electrosprays are nonpressurized, passively fed, compact and fuel-efficient electric thrusters capable of producing impulse bits in the or lower, depending on the configuration. These characteristics have the potential to allow for long-term pointing in the arcsecond range or better with practically no jitter. Experimental work is performed on a magnetically levitated testbed in vacuum to demonstrate the actuation characteristics under nonoptimal control. It is found that, subject to a noisy attitude sensor and external perturbations, electrospray thrusters are capable of producing pointing accuracies around a given set point of 22 arcseconds error during 10 h on a platform similar in size and mass to a 1U CubeSat. The implementation of such capabilities could complement or eliminate the need of reaction wheels and magnetorquers, especially in missions beyond low Earth orbit, while including propulsive capabilities for additional maneuverability in technology development or scientific missions.

Minimum sliding mode error feedback control for fault tolerant small satellite attitude control

This paper proposes a new control strategy (which we call “minimum sliding mode error feedback control, MSMEFC”) for small satellite attitude control. As we know, the attitude control algorithm plays a significant role in the whole performance of the satellite, especially under the existence of uncertain disturbances from the space. Without loss of generality, the MSMEFC is presented based on the sliding mode theory. It is assumed that the equivalent control error is defined to offset the uncertain disturbances to improve the control performance. Hence, in order to estimate the optimal equivalent control error, a cost function is derived on the basis of the principle of minimum sliding mode error. Then, the equivalent control error wills feedback to the conventional sliding mode control to obtain the final MSMEFC. According to the theoretical analyzes, the sliding mode after the MSMEFC will approximate to the ideal sliding mode, resulting in enhancing the control performance. Moreover, an adaptive non-singular terminal sliding mode is employed to compare with the performance of MSMEFC. Several simulations are performed to verify the effectiveness of proposed MSMEFC in the presence of serious perturbations, even in some fault-tolerant scenarios.

Inverse free steering law for small satellite attitude control and power tracking with VSCMGs

Recent developments in integrated power and attitude control systems (IPACSs) for small satellite, has opened a new dimension to more complex and demanding space missions. This paper presents a new inverse free steering approach for integrated power and attitude control systems using variable-speed single gimbal control moment gyroscope. The proposed inverse free steering law computes the VSCMG steering commands (gimbal rates and wheel accelerations) such that error signal (difference in command and output) in feedback loop is driven to zero. H∞ norm optimization approach is employed to synthesize the static matrix elements of steering law for a static state of VSCMG. Later these matrix elements are suitably made dynamic in order for the adaptation. In order to improve the performance of proposed steering law while passing through a singular state of CMG cluster (no torque output), the matrix element of steering law is suitably modified. Therefore, this steering law is capable of escaping internal singularities and using the full momentum capacity of CMG cluster. Finally, two numerical examples for a satellite in a low earth orbit are simulated to test the proposed steering law.

Robust Magnetic Attitude Control of Satellites

Magnetic torquers are frequently adopted as primary actuators for the attitude control of small satellites in low Earth orbit. Such actuators generate a magnetic dipole which, in turn, leads to control torques thanks to the interaction with the magnetic field of the Earth. The design of attitude control laws based on magnetic torquers is a challenging problem as the torques generated by the coils are instantaneously constrained to lie in the plane orthogonal to the local direction of the geomagnetic field vector, which varies according to the current orbital position of the spacecraft. This implies that the attitude regulation problem is formulated over a time-varying model. In this paper, the design of control laws for magnetically actuated spacecraft is considered and an approach guaranteeing robustness to parametric uncertainty and optimal performance in terms of disturbance attenuation is presented. The proposed method is based on linear time-periodic models and H∞ control theory. The results obtained by applying the proposed approach in a simulation study are also presented and discussed.

Fault tolerant small satellite attitude control using adaptive non-singular terminal sliding mode

The Attitude Control System (ACS) plays a pivotal role in the whole performance of the spacecraft on the orbit; therefore, it is vitally important to design the control system with the performance of rapid response, high control precision and insensitive to external perturbations. In the first place, this paper proposes two adaptive nonlinear control algorithms based on the sliding mode control (SMC), which are designed for small satellite attitude control system. The nonlinear dynamics describing the attitude of small satellite is considered in a circle reference orbit, and the stability of the closed-loop system in the presence of external perturbations is investigated. Then, in order to account for accidental or degradation fault in satellite actuators, the fault-tolerant control schemes are presented. Hence, two adaptive fault-tolerant control laws (continuous sliding mode control and non-singular terminal sliding mode control) are developed by adopting the nonlinear analytical model to describe the system, which can guarantee global asymptotic convergence of the attitude control error with the existence of unknown external perturbations. The nonlinear hyperplane based Terminal sliding mode is introduced into the control law design; therefore, the system convergence performance improves and the control error is convergent in “finite time”. As a result, the study on the non-singular terminal sliding mode control is the emphasis and the continuous sliding mode control is used to compare with the non-singular terminal sliding mode control. Meanwhile, an adaptive fuzzy algorithm has been proposed to suppress the chattering phenomenon. Moreover, several numerical examples are presented to demonstrate the efficacy of the proposed controllers by correcting for the external perturbations. Simulation results confirm that the suggested methodologies yield high control precision in control. In addition, actuator degradation, actuator stuck and actuator failure for a period of time are simulated to demonstrate the fault recovery capability of the fault tolerant controllers. The numerical results clearly demonstrate the good performance of the adaptive non-singular terminal control in the event of actuator fault compare with the continuous sliding mode control.

Design of Semi-Physical Simulation for Small Satellite by Virtual Display

Semi-physical simulation of attitude control system is the more synthetically test and verify for designing of small satellite control system. It is an important means of small satellite development. However, the results of current semi-physical simulation system have a lot of non-intuitive. Compare with the actual environment, the simulation environment still has striking disparity. So the shortcomings affect precision of simulation. Based on the virtual display technology, the group semi-physical simulation system has been constructed for attitude control of small satellite due to the combination with xPC real-time environment, the simulation computer, high-precision single-axis air-bearing turntable, reaction wheel, air thrust device, fiber gyroscopes, sensors synchronizer, power subsystem and wireless devices virtual display computer etc. Semi-physical simulation achieved the visual simulation in orbit and tracked new information of virtual environment of space into real-time simulation computer. Simulation results show that the simulation system for real-time attitude and orbit position of small satellite semi-physical simulation has an excellent display effect. At the same time, Real-time transfuse of orbit information provides a more accurate space environment simulation. The simulation system of small satellite attitude control to design and evaluate the more direct and convenient.

Attitude Stabilization of a Nadir-Pointing Small Satellite Using Only Magnetic Actuators

AbstractRecently, there has been a surge of interest in small satellite missions due to low-cost launch opportunities. In particular, the feasibility of nonlinear techniques for the attitude control of small satellites using only magnetic actuators is becoming a topic of increasing interest in the literature. Since magnetic control systems are relatively lightweight, require low power and are inexpensive, they are attractive for small, inexpensive satellites in low Earth orbits. In this paper we present averaging-based feedback control laws that achieve three-axis stabilized nadir-pointing attitude. We propose nonlinear quaternion feedback laws. Two cases are considered; full state feedback that uses both attitude and angular velocity measurement and passivity-based feedback that does not require angular velocity information. Computer simulations are included to illustrate the effectiveness of the proposed control laws.

A high-temperature superconductor energy-momentum control system for small satellites

A method for angular momentum management in flywheel energy storage systems that employ superconducting bearings for spacecraft attitude control is presented. Success with high-temperature superconductor momentum wheels as practical energy storage systems has led to increased interest in their dual function for providing electrical power as well as attitude control of small satellites. The physics of angular momentum management is used to quantify how superconductor magnetic bearing flywheels as electro-mechanical devices serve as both energy storage and angular momentum storage systems within an end-to-end architecture. The result is a compound energy-momentum attitude and electrical power control system for satellites.

Attitude stabilization of a satellite by magnetic coils

Stabilization problem for a satellite is considered. The only measurement is of the geomagnetic field in the satellite coordinates. The control is carried out by a magnetic moment of current coils (magnetorquers) mounted on the satellite body. The stabilizer constructed in this work solves the problems of magnetic and gravitational stabilization. Qualitative analysis and results of numerical simulation are presented. The results of simulation show that the proposed stabilization system is reliable, and has an appropriate accuracy and does not need powerful sources of energy, and therefore can be used for attitude control of small satellites.

Comparison of low-earth-orbit satellite attitude controllers submitted to controllability constraints

A rule-based fuzzy controller is presented and compared with an adaptive MIMO LQR controller in a low-Earth-orbit small satellite attitude control system. The attitude is passively gravity gradient stabilized and actively three-axis magnetorquer controlled. This method insures an Earth-pointing satellite making use of a nondepletable and nonmechanical means of control. A realistic simulation environment, using a nonlinear satellite dynamic model with linear attitude estimators plus sensor measurement noise and external disturbance torques, was used to evaluate the different control techniques.