Attitude Control Of Nanosatellites Research Articles

Design and analysis of a nanosatellite attitude control system using processor-in-the-loop approach

This paper presents an approach for the design and analysis of an attitude control system dedicated to a 3U nanosatellite using a low-cost, safe, and reliable processor-in-the-loop system for the rapid prototyping of the final product. It mainly concerns the selection between the two standard configurations of the four reaction wheels, namely the pyramid and tetrahedron configurations. The analysis was carried out by considering the presence of the mechanical uncertainties of the reaction wheels, estimated at 20% of the torque generated, and the environmental disturbances such as magnetic and gravity disturbances, solar pressure, and aerodynamic forces. It also investigates how the regulation is affected by placing the reaction wheels at or outside the spacecraft’s center of gravity. From the results obtained, it turns out that the satellite can be regulated very well regardless of the location of the reaction wheels in the satellite. When placing the actuators outside the center of gravity, a greater effort is required from the reaction wheels to change the satellite attitude, which will consume more electrical energy and wear out the wheels quickly. The torque allocation is used to distribute the forces on the four reaction wheels by means of an optimization adapted to each configuration. The analysis of degraded operation when one or two wheels fail is also studied in detail for the two mechanical configurations discussed.

Some Features of Dynamics and Attitude Control of Nanosatellites in Low Orbits

Some Features of Dynamics and Attitude Control of Nanosatellites in Low Orbits

GENETIC ALGORITHM BASED FOPID CONTROLLER FOR NANO-SATELLITE ATTITUDE CONTROL

Nano-satellites are very popular among researchers today because they are more affordable and easier to use than most large satellites. The system performance on nano-satellites needs to be improved for better by using fractional order PID (FOPID) controllers which have never been tested on unstable systems on nano-satellite objects. The PID controller development produces two fractional power parameters called the FOPID controller, which makes it even more attractive. The genetic algorithm (GA) produces the optimal computation value on the FOPID controller because it has been proven to have better performance and is improved by the ITAE performance index. Based on the analysis of responses in a steady-state in the form of overshoot, rise time and settling time on the three-axis stabilized nano-satellite attitude control, namely roll, pitch, and yaw is concluded that the FOPID controller is superior to the classic PID controller that has been previously studied. The effect of the two parameters of the FOPID controller on an unstable system for nano-satellite attitude control shows good performance results based on the ITAE performance index using the genetic algorithm (GA) method.

Experimental verification and comparison of fuzzy and PID controllers for attitude control of nanosatellites

Modern control algorithms based on artificial intelligence have shown good results in many application fields. That is why some researchers have made an effort to introduce them in the attitude control of satellites by computing their performance in simulated scenarios. However, there is a lack of experimental data due to the cost and inherent risk of testing new attitude control algorithms on board a real satellite. In this work, a laboratory nanosatellite and its testing system have been used to experimentally compare the performance of a fuzzy logic controller, a classical Proportional Integral and Derivative (PID) controller and a modified PID. Comparisons have been drawn in terms of energy consumption, convergence time, accuracy (steady state error) and robustness (with respect to variations in the external conditions). In line with previous numerical studies, the results suggest significant improvements in energy consumption, convergence time and robustness by using a fuzzy controller instead of a PID.

Fault tolerant nano-satellite attitude control by adaptive modified nonsingular fast terminal control

Fault tolerant nano-satellite attitude control by adaptive modified nonsingular fast terminal control

Impact of non-latching booms on a nanosatellite attitude control system

The Scintillation Prediction Observations Research Task (SPORT) nanosatellite is being developed in partnership with the National Aeronautics and Space Administration agency and the Brazilian Space Agency, with its launch planned for 2022. Its goal is to collect data to improve our understanding of plasma bubbles and the condition that lead to their formation, helping to predict and mitigate their interference in navigation and communication systems. Therefore, to reach this goal, the satellite has several scientific instruments to perform in-situ measurements, with five of them positioned on four booms. These booms do not have a latching system to lock their position; instead, torsional springs are employed to keep them in the deployed state, holding them against their mechanism's structure. This configuration of torsional spring and collision may lead to vibration with the potential to degrade the Attitude Determination and Control System performance, impacting the whole mission. Motivated by the lack of literature covering the non-latching booms dynamics in satellites, this work proposes a framework based on multibody dynamics to simulate satellites with such booms. It also presents a practical method to acquire experimental data and identify the parameters of the booms' deployment mechanism. Later, this work applies the proposed framework and investigates the impact of non-latching booms on the satellite control system to verify if SPORT is able to complete the maneuver. Therefore, experiments were conducted to calibrate both spring and collision models. The multibody model of the satellite was developed and later validated using commercial software. The method for capturing booms' data and determining the mechanisms' parameters shows a satisfactory performance near the booms' deployment position. The proposed framework to simulate satellites with non-latching booms is applied to the SPORT satellite. Simulations in closed-loop indicate that the booms' influence on the SPORT's control system is negligible and the satellite meets its requirements.

Parameterized fuzzy-logic controllers for the attitude control of nanosatellites in low earth orbits. A comparative studio with PID controllers

The pointing requirements for a satellite (e.g. for its payload or for its subsystems) determine the different attitude control system operational modes that are necessary along its mission. Two examples are: a low cost mode for nominal attitude stabilization; and a faster mode for antenna pointing during communication phases. In this work, an attitude control algorithm based on fuzzy logic is presented to achieve an optimal performance of the controller for different operational requirements of the satellite mission. After analysing the design process and the response of previous fuzzy controllers, several improvements have been implemented concerning the number of membership functions, their shape, and the contents of the fuzzy rules, obtaining a parameterized final fuzzy inference system whose performance depends on eighteen parameters, six per each of the three control axes. The design proposed in this work will be tested soon on board the European Space Agency OPS-SAT mission and it is easily extensible to most nanosatellite missions in Low Earth Orbits.Different sets of the eighteen tuning parameters have been used to analyse the design and the performance of this new fuzzy controller. The values have been selected through multi-objective genetic optimizations to deal with two conflicting objectives: low cost and accuracy. A numerical performance comparison shows that the new design provides more efficient controllers than those used in previous works. In addition, the performance of the new fuzzy controller has been evaluated against traditional Proportional Integrative Derivative (PID) controllers for different satellite operational requirements. The values of the PID gains have also been selected through the same multi-objective genetic optimization. The comparisons show that the new fuzzy design provides great advantages over the PID strategy in most of the operational modes considered, and how the PID strategy has difficulties to produce a low cost control law profile. Furthermore, the new fuzzy design is able to improve the Pareto front to optimal points that are unreachable with the PID schema. These results open new possibilities for the application of intelligent controllers in the space industry.

Recent development of resonant micro cavity in resonant micro-optical gyro

Resonant micro-optical gyros (RMOG) based on Sagnac effect has great potential in integration, miniaturization and sensitivity, which has broad application prospects in the fields of micro nano satellite attitude control, robot control, medical diagnosis and detection instruments, and has become a research hotspot in recent years. Resonant micro cavity as the core sensing element of resonator micro optical gyroscope, the optical characteristics of which are closely related to the performance of gyroscope system. The research progress of resonator has seriously restricted the development of RMOG. At present, integrated optical technology, micro nano optical processing technology and the application of new materials can reduce the weight and size of the resonator, reduce the cost and power consumption, and increase the reliability and performance index of the system. Combined with the information disclosed by periodical conference and related research institutions, the development status, basic principle and characteristic parameters of resonant micro optical gyroscope were briefly introduced. Various new structure designs of resonant micro cavity at home and abroad were listed, and the characteristics and potential of different structures were analyzed. In addition, the new materials for fabrication of resonant micro cavity at home and abroad were reviewed, and the optical properties of different materials were summarized. The future development direction and technical development path of resonator micro cavity for resonator micro optical gyroscope were preliminarily discussed.

Attitude control of nanosatellite with single thruster using relative displacements of movable unit

The attitude dynamics of a nanosatellite (NS) with one movable unit changing its angular position relative to a main body of the nanosatellite is considered. This relative movability of the unit can be implemented with the help of flexible rods of variable length connecting the unit with the main body. Change of the relative position of the movable unit shifts the center of mass of the entire mechanical system. The NS has a single jet engine rigidly mounted into the NS main body. The shift of the mass center creates an arm of the jet-engine thrust and a corresponding control torque. Schemes to control the attitude dynamics of the satellite using the movability of its unit are developed, using both the torque from the engine and inertia change.

Generic Model of a Satellite Attitude Control System

A generic model of a nanosatellite attitude control and stabilization system was developed on the basis of magnetorquers and reaction wheels, which are controlled by PID controllers with selectable gains. This approach allows using the same architectures of control algorithms (and software) for several satellites and adjusting them to a particular mission by parameter variation. The approach is illustrated by controlling a satellite attitude in three modes of operation: detumbling after separation from the launcher, nominal operation when the satellite attitude is subjected to small or moderate disturbances, and momentum unloading after any reaction wheel saturation. The generic control algorithms adjusted to each mode of operation were implemented in a complete attitude control system. The control system model was embedded into a comprehensive simulation model of satellite flight. The simulation results proved the efficiency of the proposed approach.

Design of High Pointing Accuracy NPSAT-1 Satellite Attitude Systems of Armature Controlled DC Motor with utilization for PD Controller

An Attitude control system plays the important role to maintain the satellite to desired attitude orientations. The intended application of NANO satellite in low earth orbits (LEO) helps find transient responses with and without controllers. LEO satellites typically orbit at an altitude ranging between160-2000 km. LEO satellites are widely used for remote sensing, navigation, and military surveillance applications. The Nano NPSAT-1 satellite attitude control systems (ACS) are described in this research work. The high pointing accuracy attitude estimation and feedback control systems are presented. The design specifications have been taken to meet the accuracy requirements (desired value ≤ 0.2 seconds) of Nano satellite attitude control. The feedback signal from on-board sensors compared with reference orbit trajectory and implementation of the Proportional Derivative (PD) controller is constructed. An algorithm of Nano satellite (NPSAT-1) attitude control is implemented using MATLAB Tools. In addition, the closed loop poles help find the gain of the system using Root Locus (RL) methods. The satellite control system is used to improve the transient response like overshoot and settling time of the system. Thus, the design of attitude control to improve the rise time, the settling time, the maximum overshoot, and no steady state error was carried out.

Development of Fault Detection and Identification Algorithm Using Deep learning for Nanosatellite Attitude Control System

Satellites should have high reliability because they are required to operate autonomously, while performing a given mission. To realize this, it is necessary to use heritage parts or include redundancy. However, for nanosatellites, where it is difficult to include redundancy due to volume and weight limitations, fault management becomes crucial. In this study, a new method based on deep learning is proposed for detecting and identifying the faults in the reaction wheel, which is one of the satellite actuators. A deep learning model is applied to learn the fault and fault type using the residual between the measured attitude data and the estimated attitude date. In this study, it is assumed that three reaction wheels are installed, and fault detection is designed accordingly. The proposed model enables the satellite to detect faults autonomously, even when it is not communicating with the ground station, and is expected to be highly beneficial for the autonomous operation of the mega-constellation mission using nanosatellite that is to be activated in future.

Intelligent Fuzzy PD+I Controller with Stabilizer for Nano Satellite Attitude Control System

This paper discusses the variation of proportional and derivative gains in a system under the control of PD+I controller that uses fuzzy inference method an effort is made to design the fuzzy proportional-derivative (PD) + I controller for a nonlinear the proposed fuzzy PD + I structure shows superior performance in the Nano-satellite attitude control system. Given that 2-input PD type controllers are commonly used to control Nano satellite attitude systems (Innovative Satellite InnoSAT), This paper analysis the performance of Fuzzy PD+I on satellite system. Fuzzy PD+I controller design can control Nano satellite attitude control system nicely and utilized the Fuzzy PD+I controller which optimize the stability of the Nano Satellite control system. In addition, the scheme is robust to variations in nonlinearities, as well as the steady-state gain of the plant. We illustrate the effectiveness of our scheme In MATLAB / SMULINK simulation environment.

A photodiode based miniature sun sensor

The solar vector is one of the most important parameters for attitude control of nanosatellites. This attitude control must be achieved without the sensors adding significantly to its size or mass. This paper presents a photodiode-based miniature sun sensor, which consists of two triangular pyramidal sensor unit structures, with each unit comprising three micro-silicon photodiodes. The two sensor units are installed on the diagonal of the nanosatellite to form a complete sun sensor capable of achieving a full-field range of solar vector measurements. In this paper, the mathematical model of the short-circuit currents of the silicon photodiodes as a function of the solar vector coordinates is deduced. A sensor sample was built and installed on a nanosatellite model, and the temperature compensation coefficient of the silicon photodiodes was obtained experimentally. The dynamic characteristic, linearity, hysteresis and repeatability of the component were measured. The sun sensor introduced in this paper can be placed on any satellite platform to allow a full range solar vector measurement, and this would result in an increase of only 1.86 g and 0.9 cm3 of the satellite’s mass and volume, respectively.

Mathematical model of angular motion of nanosatellites with inertial actuators

The problem of developing a nanosatellite attitude control system using three reaction wheels mounted along the main central axes of inertia is discussed in the paper. The law of controlling a nanosatellites attitude is based on a PD - controller. The stability of the process of controlling the nanosatellite attitude using the Lyapunov function method has been analyzed. It allows us to prove that the obtained control law provides the asymptotic stability of nanosatellite angular motion. Hereafter, the function of control voltage for electric motors of reaction wheels taking into account their specifications is obtained based on the developed mathematical model of reaction wheel dynamics. Numerical calculations of controlled angular motion have been carried out for the nanosatellite CubeSat3U with actuators on the basis of a commercial DC motor. Several cases of controlling the satellites rotation by different angles are considered in the course of the numerical experiments. The results of numerical experiments showed the adequacy of the developed mathematical model.

In-orbit offline estimation of the residual magnetic dipole biases of the POPSAT-HIP1 nanosatellite

The nanosatellite POPSAT-HIP1 is a Cubesat-class spacecraft launched on the 19th of June 2014 to test cold-gas based micro-thrusters; it is, as of April 2015, in a low Earth orbit at around 600km of altitude and is equipped, notably, with a magnetometer. In order to increment the performance of the attitude control of nanosatellites like POPSAT, it is extremely useful to determine the main biases that act on the magnetometer while in orbit, for example those generated by the residual magnetic moment of the satellite itself and those originating from the transmitter.Thus, we present a methodology to perform an in-orbit offline estimation of the magnetometer bias caused by the residual magnetic moment of the satellite (we refer to this as the residual magnetic dipole bias, or RMDB). The method is based on a genetic algorithm coupled with a simplex algorithm, and provides the bias RMDB vector as output, requiring solely the magnetometer readings. This is exploited to compute the transmitter magnetic dipole bias (TMDB), by comparing the computed RMDB with the transmitter operating and idling.An experimental investigation is carried out by acquiring the magnetometer outputs in different phases of the spacecraft life (stabilized, maneuvering, free tumble). Results show remarkable accuracy with an RMDB orientation error between 3.6° and 6.2°, and a module error around 7%. TMDB values show similar coherence values.Finally, we note some drawbacks of the methodologies, as well as some possible improvements, e.g. precise transmitter operations logging. In general, however, the methodology proves to be quite effective even with sparse and noisy data, and promises to be incisive in the improvement of attitude control systems.

Dynamics of Spring-Deployed Solar Panels for Agile Nanospacecraft

Advanced performances in nanospacecraft are typically associated with high onboard power availability, requiring deployed solar arrays. Nanospacecraft geometrical and weight constraints do not allow the use of reversible dynamics and lock systems used in conventional spacecraft. Most nanospacecraft deployed solar panel systems are therefore based on compact and reliable miniaturized mechanisms, such as simple passively deployed preloaded springs. The use of these kinds of deployed solar arrays may interfere with the performance of the attitude control system. Agile nanospacecraft, designed to perform rapid attitude maneuvers, may potentially be affected by the solar array dynamical interaction with the main satellite body motion. In addition, the presence of inherently nonlinear elements, such as bounds on allowed rotations and preloaded springs, deserve a dedicated analysis. The analysis developed in this paper is based on the performance of state-of-the art agile nanosatellite attitude control systems, specifically developed for delivering high torque and fast attitude maneuvering, such as miniaturized control moment gyros. By numerical simulation, it is shown that, despite the simplicity of the deployment system and the absence of locking mechanisms, the attitude dynamics remains substantially unaffected by the solar array dynamics. This is obtained straightforwardly by properly sizing the spring stiffness and by a sufficient level of preload.

Design of Attitude Control Systems for CubeSat-Class Nanosatellite

We present a satellite attitude control system design using low-cost hardware and software for a 1U CubeSat. The attitude control system architecture is a crucial subsystem for any satellite mission since precise pointing is often required to meet mission objectives. The accuracy and precision requirements are even more challenging for small satellites where limited volume, mass, and power are available for the attitude control system hardware. In this proposed embedded attitude control system design for a 1U CubeSat, pointing is obtained through a two-stage approach involving coarse and fine control modes. Fine control is achieved through the use of three reaction wheels or three magnetorquers and one reaction wheel along the pitch axis. Significant design work has been conducted to realize the proposed architecture. In this paper, we present an overview of the embedded attitude control system design; the verification results from numerical simulation studies to demonstrate the performance of a CubeSat-class nanosatellite; and a series of air-bearing verification tests on nanosatellite attitude control system hardware that compares the performance of the proposed nonlinear controller with a proportional-integral-derivative controller.

Attitude Control of Nanosatellites by Paddle Motion Using Elastic Hinges Actuated by Shape Memory Alloy

Recent research has been extending the applications of small satellites called microsatellites, nanosatellites, or picosatellites. To further improve capability of those satellites, a lightweight, active attitude-control mechanism is needed. This paper proposes a concept of inertial orientation control, an attitude control method using movable solar arrays. This method is made suitable for nanosatellites by the use of shape memory alloy (SMA)-actuated elastic hinges and a simple maneuver generation algorithm. The combination of SMA and an elastic hinge allows the hinge to remain lightweight and free of frictional or rolling contacts. Changes in the shrinking and stretching speeds of the SMA were measured in a vacuum chamber. The proposed algorithm constructs a maneuver to achieve arbitrary attitude change by repeating simple maneuvers called unit maneuvers. Provided with three types of unit maneuvers, each degree of freedom of the satellite can be controlled independently. Such construction requires only simple calculations, making it a practical algorithm for a nanosatellite with limited computational capability. In addition, power generation variation caused by maneuvers was analyzed to confirm that a maneuver from any initial attitude to an attitude facing the sun was justifiable in terms of the power budget.

Laboratory tutorial practice with facility for attitude control simulation

The paper presents a combined approach to teaching of Spaceflight Dynamics, Theoretical Mechanics and Control Theory. In addition to a standard lecture course, this approach provides hands-on training at the laboratory, offering to the student a possibility to develop, design, assemble and test his/her own project of a nanosatellite attitude control system. We describe the software for mathematical modeling, the facilities for laboratory simulation of controlled system dynamics and student training based on these facilities.