Performance Of Attitude Control System Research Articles

Research on UAV flight attitude control system based on simulink and genetic algorithm

With the evolution of networking and information technology, unmanned aerial vehicles (UAVs) have emerged as integral components in numerous fields, executing diverse and vital tasks. The accuracy and efficiency of UAV operations are greatly influenced by their attitude control systems, which dictate flight stability, speed, and functionality. However, traditional control methods for UAV flight stability and accuracy have become outdated, unable to meet the demands of modern applications. To address these challenges, this study introduces a novel attitude control approach based on genetic algorithms. This method utilizes the principles of natural selection and genetic traits to optimize the parameters of the attitude control system. The objective is to enhance the current control effectiveness, improve the UAV's maneuverability, and enable its versatile applications across multiple fields. Simulink tool simulations are conducted to evaluate the stability of the optimized results. This simulation tool allows for the replication of real-world flight conditions and the assessment of the attitude control system's performance under various scenarios. The results of these simulations provide valuable insights into the effectiveness of the genetic algorithm-based attitude control method. This study contributes to the advancement of UAV technology by improving the accuracy and stability of attitude control systems. It paves the way for more efficient and reliable UAV operations in various fields, from environmental monitoring to search-and-rescue missions, thus enhancing their operational capabilities and expanding their potential applications.

A novel output redefinition method for nonlinear H∞ attitude stabilization of a flexible spacecraft

AbstractThis paper addresses a novel output redefinition method for the nonlinear attitude stabilization problem of a flexible spacecraft. As an idea, the flexible appendage tip‐point can be selected as attitude output which may improve attitude pointing accuracy. However, it makes the system non‐minimum phase. The output redefinition method is an effective method to make the system minimum phase. In this paper, a novel output redefinition is proposed to overcome the non‐minimum phase property such that it improves the attitude control system performance. Then, a nonlinear controller is designed based on back‐stepping approach and a modal observer. The simulation results show the effectiveness of the proposed method in the presence of model uncertainties and environmental disturbances.

Nonlinear finite time attitude control of flexible spacecraft based on a novel output redefinition method

A finite time attitude controller is designed for a flexible spacecraft based on a novel output redefinition method, in this paper. To make the flexible appendages vibration suppression effective, the appendage tip-point is selected as the output. First, a novel output redefinition method is proposed to overcome the non-minimum phase property of the dynamic model. The proposed method not only makes the system model minimum phase but also improves the attitude control system performance. Consequently, the precise attitude pointing and stabilization are achieved. Then, a nonlinear finite time H∞ controller is designed based on the backstepping approach. For the situation where the modal variables measurements are not available, a modal observer is also designed. The simulation results show the effectiveness of the proposed method in the presence of the model uncertainties and environmental disturbances.

Statistical Investigation about the SDRE Optimality for Satellites with Nonlinearities

The Attitude Control System (ACS) of a Satellite can be designed with success by linear control theory if the satellite has slow angular motions. However, for fast maneuvers, the linearized models are not able to represent the effects of the nonlinear terms. One candidate technique to design the ACS of the satellite under fast maneuvers is the State-Dependent Riccati Equation (SDRE), which provides an effective algorithm for synthesizing nonlinear feedback control taking into account the nonlinearities. Nonetheless, much criticism has been leveled against the SDRE method since it does not assure global asymptotic stability (GAS), because there are situations in which global asymptotic stability cannot be achieved (e.g., systems with multiple equilibrium points). One way to study the GAS is by estimating the region of attraction (ROA) which in turn is fundamental to investigate the performance of the controller designed by the SDRE method The Brazilian National Institute for Space Research (INPE) is responsible to build a Remote-Sensing Satellites, called Amazonia-1. The ACS of the Amazonia-1 satellite must be stabilized in three axes so that the optical payload can point to the desired target. In this paper, one investigates the performance of the Amazonia-1 ACS using a statistical test (unpaired t-test) that compares the optimality inside the ROA controlled by LQR and SDRE. By several simulations running a full Monte Carlo perturbation satellite model one observes a significant difference between the optimality of the two controllers in favor of SDRE when nonlinearities are accounted for.

H∞ Control Design for a Rigid-Flexible Spacecraft under a Fuel Slosh Influence

The spacecraft Attitude Control System (ACS) performance and robustness depend on the interactioneffects between the fuel slosh motion, the panel's flexible motion, and the spacecraft rigid motion, mainly duringtranslational and/or rotational maneuvers. In regards to satellite pointing accuracy flexibility and fuel, slosh is thetwo most important effects that should be considered in the satellite ACS design since their interactions can damage the ACS performance and robustness. Once, the lowest vibration frequencies, normally of the sloshing modeare about six times less than of the ACS bandwidth. Therefore, there is a strong possibility that this mode can destabilize the ACS pointing accuracy. This phenomenon is called spillover because the control effort spills over outside the control bandwidth. As a result, the designer needs to explore the limits between the conflicting requirements of performance, that is, increase of the bandwidth without introduction noise in the ACS keeping the systemrobustness to parameters variation. In this paper, one applies the H infinity control method which can deal withthese two design requirements (performance and robustness) considering the controller error pointing that may belimited by the minimum time necessary to suppress disturbances that affect the satellite attitude acquisition. Theequations of motions are obtained considering the Lagrange method for small flexible deformations and a mechanical model of liquid sloshing which allows modeling and investigating the longitudinal dynamic characteristics of apartially filled liquid tank during a pitch maneuver, satisfying performance and robustness requirements.

Cassini Spacecraft Attitude Control System Performance and Lessons Learned, 1997–2017

A sophisticated interplanetary spacecraft, Cassini–Huygens was launched on 15 October 1997. After achieving orbit at Saturn in 2004, Cassini collected science data throughout its four-year prime mission (2004–2008) and nine-year extended mission (2008–2017). The Cassini Attitude and Articulation Control Subsystem (AACS) is perhaps the spacecraft subsystem that must satisfy the most mission and science pointing requirements. The performance of the Cassini AACS design was superb from launch through end of mission. All key mission and science requirements were met with significant margins. Overviews of the Cassini attitude control system design and flight performance (in 1997–2017) are given in this paper. Processes taken by the attitude control operations team to guard against human errors are also outlined. Cassini flight experience and many lessons learned are equally applicable to the safe operations of other interplanetary missions.

Discussions on attitude determination and control system for micro/nano/pico-satellites considering survivability based on Hodoyoshi-3 and 4 experiences

The recent advancement of micro/nano/pico-satellites technologies encourages many universities to develop three axis stabilized satellites. As three axis stabilization is high level technology requiring the proper functioning of various sensors, actuators and control software, many early satellites failed in their initial operation phase because of shortage of solar power generation or inability to realize the initial step of missions because of unexpected attitude control system performance. These results come from failure to design the satellite attitude determination and control system (ADCS) appropriately and not considering “satellite survivability.” ADCS should be designed such that even if some sensors or actuators cannot work as expected, the satellite can survive and carry out some of its missions, even if not full. This paper discusses how to realize ADCS while taking satellite survivability into account, based on our experiences of design and in-orbit operations of Hodoyoshi-3 and 4 satellites launched in 2014, which suffered from various component anomalies but could complete their missions.

Vibration suppression-based attitude control for flexible spacecraft

In this paper, a novel approach for high-accuracy attitude control in a flexible spacecraft is developed by compensating and damping the vibration of the flexible appendage. A vibration suppression-based attitude controller (VSAC) that addresses the coupling disturbance in attitude dynamics that is generated by the vibration of the flexible appendage is developed for flexible spacecraft. The proposed VSAC integrates the existing attitude control theories with vibration suppression control strategies by combining a time delay feedback controller with a PD controller. The stability of the flexible spacecraft control system under the proposed VSAC is investigated in time domain, and the damping effect of the proposed VSAC on the vibration of the flexible appendage is analyzed in frequency domain. The proposed VSAC utilizes the traditional attitude controller and vibration suppression controller, which can not only effectively improve the performance of attitude control system but also efficiently damp the vibration of the flexible appendage. Finally, numerical simulations are performed, and analysis and conclusions are presented.

SATELLITE ATTITUDE CONTROL SYSTEM DESIGN WITH NONLINEAR DYNAMICS AND KINEMTICS OF QUATERNION USING REACTION WHEELS

The success of a space mission where the satellite must perform rapid attitude maneuvers with great angles is extremely dependent of a careful investigation of the nonlinear dynamics of the satellite. Since these big maneuvers imply in the dynamic coupling between the satellites angular motion and the actuators such as reaction wheels and/or gas jets. As a result, this coupling must be taking into account in the Attitude Control System (ACS) design. This paper presents the nonlinear model derivation of a rigid satellite and the performance comparison of two controllers designed by Lyapunov and LQR methods. The dynamics of the satellite is initially derived in the non-linear Euler equations form and the kinematics is based on the quaternion parametrization which represent the rotation and attitude motion, respectively. In the sequel, the linear model is obtained where linearization is about an operating point of the arbitrary angular velocity and the reaction wheel angular momentum. From this model, several simulations are performed in order to investigate the influence of the nonlinear dynamics in the in the SCA performance which is designed by trial and error and by the Linear Quadratic Regulator approaches. The ACS performance is evaluated considering the capacity of the reaction wheels to maintain the stability and to control the angular velocity and the attitude of the satellite. The stability is investigated comparing the location of the poles and zeros of the open and closed loops. The ACS performance is evaluated comparing the amount of energy spend by each control law. Keywords: Satellite Control, nonlinear dynamics, LQR and Lyapunov theory

Bounded finite-time attitude tracking control for rigid spacecraft via output feedback

This paper investigates a velocity-free finite-time attitude control scheme for a rigid spacecraft with actuator saturation and external disturbances. Initially, a finite-time observer is designed to compensate the unknown angular velocity information. With the estimated values, a finite-time controller is further developed under which the time-varying reference attitude trajectory can be tracked precisely. Moreover, input saturation problem is solved by adaptive method. Rigorous Lyapunov-based analysis shows that the states of the closed-loop system converge to a small neighborhood of the origin in finite time. Finally, the dynamic performance of attitude control system is presented by numerical simulation examples and the superiority of the finite-time control scheme is verified.

Vibration Suppression of an Underactuated Dynamic System Using Virtual Actuators

Sloshing is one of the critical problems in aerospace vehicles with liquid containers. Motion of the liquid in resonance situations can degrade the stability and performance of attitude control systems. Two important characteristics of this time varying phenomenon are sensorless and underactuated properties which lead to difficulty of attitude control system design. In this paper, a new technique based on soft sensor and virtual actuator is used to suppress the effects of fuel sloshing in an aerospace launch vehicle (ALV). For this purpose, a nonlinear dynamic model of the vehicle with mechanical model of the fuel sloshing is considered as a multibody dynamic system. The preliminary attitude control system of the vehicle is extended using the new vibration suppression technique and a numerical simulation of the nonlinear model is carried out. Results of the simulation show that the undesired effects of the fuel sloshing are effectively decreased using the proposed vibration suppression technique.

Stabilization and parameter identification of tumbling space debris with bounded torque in postcapture

This paper presents a new control scheme for the problem of a space robot after capturing an unknown tumbling target, such as space debris. Robotic capturing the target may destabilize the base of spacecraft and control torque is bounded which would affect the performance of attitude control system. To stabilize the base with bounded torque in postcapture scenario, a new control scheme which utilizes the control torque to balance angular momentum and motion of the manipulator to compensate limitation of the torque, is proposed. Considering uncertainties of the target, parameter identification technique for tumbling target with linear momentum is utilized to correct parameters of the controller. To verify validity and feasibility of the proposed concept, a planar space robot capturing small, medium and large target with or without linear momentum is studied. The results show that the whole system is stabilized finally and all the inertial parameters of the target converge to their real values.

Design of Satellite Attitude Control System Considering the Interaction between Fuel Slosh and Flexible Dynamics during the System Parameters Estimation

The design of the satellite Attitude Control System (ACS) becomes more complex when the satellite structure has different type of components like, flexible solar panels, antennas, mechanical manipulators and tanks with fuel, since the ACS performance and robustness will depend if the dynamics interaction effects between these components are considered in the satellite controller design. A crucial interaction can occur between the fuel slosh motion and the satellite rigid motion during translational and/or rotational maneuver since these interactions can change the satellite center of mass position damaging the ACS pointing accuracy. Although, a well-designed controller can suppress such disturbances quickly, the controller error pointing may be limited by the minimum time necessary to suppress such disturbances affecting thus the satellite attitude acquisition. It is known that one way to minimize such problems is to design controllers with a bandwidth below the lowest slosh and/orvibration mode which can result in slow maneuvers inconsistent with the space mission requirements. As a result, the design of the satellite controller needs to explore the limits between the conflicting requirements of performance and robustness. This paper investigates the effects of the interaction between the liquid motion (slosh) and the flexible satellite dynamics in order to predict what the damage to the controller performance and robustness is. The fuel slosh dynamics is modeled using its pendulum analogs mechanical system which parameters are identified using the Kalman filter technique. This information is used to designs and to compare the satellite attitude control system by the Linear Quadratic Regulator (LQR) and the Linear Quadratic Gaussian (LQG methods. Besides, one investigates the effects of the rod length estimation in the plant of the system stability. This investigation has shown that the poles of the plant to walk to and from the imaginary axis, leaving in the end the plant more stable.

Robust Fault-Tolerant Control for Satellite Attitude Stabilization Based on Active Disturbance Rejection Approach with Artificial Bee Colony Algorithm

This paper proposed a robust fault-tolerant control algorithm for satellite stabilization based on active disturbance rejection approach with artificial bee colony algorithm. The actuating mechanism of attitude control system consists of three working reaction flywheels and one spare reaction flywheel. The speed measurement of reaction flywheel is adopted for fault detection. If any reaction flywheel fault is detected, the corresponding fault flywheel is isolated and the spare reaction flywheel is activated to counteract the fault effect and ensure that the satellite is working safely and reliably. The active disturbance rejection approach is employed to design the controller, which handles input information with tracking differentiator, estimates system uncertainties with extended state observer, and generates control variables by state feedback and compensation. The designed active disturbance rejection controller is robust to both internal dynamics and external disturbances. The bandwidth parameter of extended state observer is optimized by the artificial bee colony algorithm so as to improve the performance of attitude control system. A series of simulation experiment results demonstrate the performance superiorities of the proposed robust fault-tolerant control algorithm.

Design of Attitude Control System for UAV Based on Feedback Linearization and Adaptive Control

Attitude dynamic model of unmanned aerial vehicles (UAVs) is multi-input multioutput (MIMO), strong coupling, and nonlinear. Model uncertainties and external gust disturbances should be considered during designing the attitude control system for UAVs. In this paper, feedback linearization and model reference adaptive control (MRAC) are integrated to design the attitude control system for a fixed wing UAV. First of all, the complicated attitude dynamic model is decoupled into three single-input single-output (SISO) channels by input-output feedback linearization. Secondly, the reference models are determined, respectively, according to the performance indexes of each channel. Subsequently, the adaptive control law is obtained using MRAC theory. In order to demonstrate the performance of attitude control system, the adaptive control law and the proportional-integral-derivative (PID) control law are, respectively, used in the coupling nonlinear simulation model. Simulation results indicate that the system performance indexes including maximum overshoot, settling time (2% error range), and rise time obtained by MRAC are better than those by PID. Moreover, MRAC system has stronger robustness with respect to the model uncertainties and gust disturbance.

Satellite Attitude Control System Design Taking into Account the Fuel Slosh and Flexible Dynamics

The design of the spacecraft Attitude Control System (ACS) becomes more complex when the spacecraft has different type of components like, flexible solar panels, antennas, mechanical manipulators and tanks with fuel. The interaction between the fuel slosh motion, the panel’s flexible motion and the satellite rigid motion during translational and/or rotational manoeuvre can change the spacecraft center of mass position damaging the ACS pointing accuracy. This type of problem can be considered as a Fluid-Structure Interaction (FSI) where some movable or deformable structure interacts with an internal fluid. This paper develops a mathematical model for a rigid-flexible satellite with tank with fuel. The slosh dynamics is modelled using a common pendulum model and it is considered to be unactuated. The control inputs are defined by a transverse body fixed force and a moment about the centre of mass. A comparative investigation designing the satellite ACS by the Linear Quadratic Regulator (LQR) and Linear Quadratic Gaussian (LQG) methods is done. One has obtained a significant improvement in the satellite ACS performance and robustness of what has been done previously, since it controls the rigid-flexible satellite and the fuel slosh motion, simultaneously.

An anti-disturbance PD control scheme for attitude control and stabilization of flexible spacecrafts

This paper studies the attitude control problem of spacecrafts with flexible appendages. It is well known that the unwanted vibration modes, model uncertainty and space environmental disturbances may cause degradation of the performance of attitude control systems for a flexible spacecraft. In this paper, the vibration from flexible appendages is modeled as a derivative-bounded disturbance to the attitude control system of the rigid hub. A disturbance-observer-based control (DOBC) is formulated for feedforward compensation of the elastic vibration. The model uncertainty and space environmental disturbances as well as other noises are merged into an “equivalent” disturbance. We design a composite controller with a hierarchical architecture by combining DOBC and PD control, where DOBC is used to reject the vibration effect from the flexible appendages. Numerical simulations are performed to demonstrate that by using the composite hierarchical control law, disturbances can be effectively attenuated and the robust dynamic performances be enhanced.

A new method for performing digital control system attitude computations using quaternions

Space vehicle attitude control system performance is often limited by the computational speed of airborne digital computer hardware. The application of quaternions to digital atti- tude control problems involving coordinate system transformations can reduce computation time by more than 40% over the equivalent direction cosine matrix solution due to a new method of treating quaternions which allows the order of multiplication to be interchanged so as to isolate the most rapidly varying parameters. This paper presents a formulation for control equations in which the control error is expressed as a and all coordinate system transformations are performed using quaternions. The new principle of quaternion algebra which allows interchanging the order of multiplicatio n is developed and used to simplify the control equations. The control equations are applied to a three- gimbal IMU example and the results are compared to an equivalent direction cosine matrix solution. T HE problem of describing the relationship between two coordinate systems is one of the most basic concepts en- countered in the field of navigation and guidance. Until re- cently Euler angles formed the most widely used method of describing the rotation between two coordinate systems. The use of three Euler angles to fix the attitude of a body with re- spect to an inertial or reference coordinate system has the advantage of being well-defined geometrically and fairly simple to visualize. The 3X3 direction cosine matrix, how- ever, was more readily adaptable to high-speed digital com- putation and has replaced the Euler angle method in the solution of all but the most simple navigation problems. The direction cosine matrix approach is particularly useful to de- scribe several successive rotations of a body with respect to a fixed reference system. A third, but infrequently applied, approach to establishing body orientation utilizes the qua- ternion, first devised by Hamilton. The ap- proach makes use of Euler's Theorem which states that any real rotation of one coordinate system with respect to another may be described by a rotation through some angle about a single fixed axis. The is a compact form for rep- resenting the single fixed axis and angle referred to by Euler's Theorem. The may be handled much the same as the direction cosine matrix, in that successive rotations result in successive multiplication. The ad- vantage of the lies in its ability to define the ro- tational relationship between two coordinate systems using only four numbers as opposed to the nine elements of a direc- tion cosine matrix. This results in a similar simplification when the effect of several successive rotations is being com- puted. The principal impediment to use of quaternions has been that the direction cosine matrix, rather than the qua- ternion, is the desired end product of computation. Usually, the computation saved by using quaterion multiplication to perform successive coordinate system rotations is lost when the resultant must be put in the form of a direction cosine matrix. When the object of the computation is to obtain the control error it is not necessary to define the direction cosine matrix explicitly. In fact, optimum control, in the least angle sense, is performed by defining the control error in terms of the single axis and angle of Euler's Theorem. This may be done by formulating the problem completely in terms of quater- nions. The formulation is shown to be particu- larly advantageous due to a new method for treating quater- nions which allows the order of multiplicatio n to be inter- changed.