Zero-Speed Crossing Avoidance with Three Active Reaction Wheels using Set-Point Angular Momentum Management

High-resolution satellite images (HRSIs) obtained from linear array charge-coupled device sensors always suffer from geometric instability in the presence of attitude jitter. Therefore, detection and compensation of spacecraft attitude jitter in both the cross-track and along-track directions are crucial to improve the geometric accuracy of HRSIs. A number of reports have been made on the detection and estimation of cross-track attitude jitter. However, the detection of the attitude jitter in the along-track direction is more complicated due to the impact of topographic change. This paper presents a novel approach to achieve accurate estimation of the along-track attitude jitter by eliminating the influence of topographic information based on the back-projection residuals of three-line-array (TLA) images. The principle of detection and estimation of along-track attitude jitter is described, and the proposed approach consists of three main components as follows: 1) dense image matching of the TLA images using a comprehensive matching strategy; 2) detection of the back-projection residuals in the line direction caused by attitude jitter; and 3) estimation of the along-track attitude jitter from the back-projection residuals using a genetic algorithm. Experiments were conducted using China's Ziyuan-3 (ZY-3) TLA images, and the experimental results reveal that the frequency of the attitude jitter in the along-track direction ranges between 0.6 and 0.7 Hz, which is consistent with the frequency in the cross-track direction observed in our previous study. In addition, a comparison of the results of the proposed approach with those from direct attitude observations shows good consistency, with as little as 0.1-pixel disparity, which demonstrates the feasibility and reliability of the proposed approach. Furthermore, the geometric accuracy is further improved from a pixel level to a subpixel level and the periodic trend is removed with the compensation of the estimated attitude jitter in addition to the conventional affine compensation, which validates the potential of the proposed approach for geometric accuracy improvement with ZY-3 TLA images.

The problem of optimal synthesis of a nonlinear system’s control law parameters of satellite attitude control under constraints is considered. The maximum stability degree of a system is considered as an optimality criterion. The reaction wheels’ control moments and angular velocities achievable within the drives’ physical characteristics are considered as constraints. A linear model of the system converted from the original nonlinear model was used to solve the problem. The decomposition method, based on the transition from real time to relative time, is used to separate the tasks of optimality criterion maximization and constraints consideration. A method of synthesizing the optimal form of system transient process according to the maximum stability degree criterion is developed. An analytical method for determining the optimal values of control law parameters is proposed, based on the structural features of the linear differential equations system. A method that takes into account constraints on control moments and angular velocities of reaction wheels, based on transition scale from relative time to real time and its calculation algorithm, ensuring that conditions of all constraints are met, is developed. An example of solving the problem of optimal system synthesis is considered. The results demonstrate the effectiveness and simplicity of the proposed methods and algorithm for targeted selection of a system’s technical characteristics, taking into account constraints on reaction wheels’ maximum angular velocities and control moments values.

This paper presents a control allocation method for simultaneous station-keeping and momentum management maneuvers. The considered satellite model is equipped with internal and external actuators to control the satellite’s orbit and attitude and, concurrently, unload the angular momentum stored in the reaction wheels. Our method allocates the external and internal actuators to achieve the control objectives. Internal actuators such as reaction wheels generate torque based on the conservation of the momentum. In this study, two different controllers are utilized to control satellite attitude and reaction wheel speeds. To manage the satellite attitude in three axes, at least three reaction wheels are needed. In this study, four reaction wheels are used to ensure control in case of failure of a reaction wheel. In addition, with six chemical thrusters, east-west and north-south orbit correction maneuvers are performed. For the satellite to serve throughout its service life, fuel optimization of the satellite is required. The proposed control allocation method will enforce constraints that maintain orbital accuracy and the satellite in a nadir pointing attitude configuration while minimizing the use of thrusters, significantly reducing fuel consumption. The method combines the two generally separated objectives of orbital and attitude control through constraints determined by the propulsion system. Numerical simulations are performed to validate the proposed method. The simulations show that using the proposed control allocation method can significantly increase the service life of geostationary satellites.

Reaction wheels are used as actuators for attitude control of satellites. Ball bearing is being used in reaction wheel application since many decades. Even small variations in friction torque in ball bearings affects the pointing stability of the satellites used for high resolution imaging application. Also the bi-directional operation of reaction wheel requires frequent zero speed crossing. The stiction present in ball bearing wheels affects the stability during zero speed cross over. Magnetic Bearings have distinct advantages over conventional bearings, like zero friction (due to non-contact operation), no lubrication, long life, less power, etc. Absence of lubrication in magnetic bearing makes it compatible to harsh environments like vacuum. It provides smooth zero speed cross over due to negligible stiction, which is advantageous to achieve better pointing stability. The context of this paper benchmarks the design, simulation and test results of a two axes actively controlled magnetic bearing. The wheel is designed to have adequate slew rate capability to ensure non-contact operation during satellite maneuvers.

Cost reduction on system tests with hardware in the loop during qualification phase of the onboard attitude control software imply in increasingly simulation realism. Attitude simulation software not only has to mimic the dynamical behavior of the spacecraft, but also shall be easily configurable, to adapt to different orbit, attitude pointing requirements, environment disturbances, sensors, actuators, and to run in real-time among other minor requirements. As the requirements for accuracy and stability for attitude pointing has increased in many modern missions, so do the complexity of the simulation algorithms and necessity for dynamical realism. Rigid body dynamics can be considered excessively simple to model a satellite with several reaction wheels, non-rigid structures, unbalanced solar arrays, fuel and liquid sloshing, not to mention crew motion and vehicle docking. This work presents the efforts being carried out to fulfill the requirements of the Brazilian space missions, nominally Multi Mission Platform, with a simulation environment capable to accomplish from early definition and mission analysis phase up to acceptance tests of the attitude control software. Development is being conducted in order to achieve a high degree of realism and range of applicability of the simulation software, compatible with the expected cost reduction of the simulation hardware (ordinary PCs). The kinematics and dynamic models to support AOCS (Attitude and Orbit Control Subsystem) simulation and testing with or without hardware-in-the-loop for Brazilian space missions are presented. Normally the dynamic equations are expressed in angular momentum components, due to the high complexity of these equations when derived in terms of the angular velocity. This, of course, is particularly important in satellites composed by several articulated rigid bodies as solar panels, robotic arms, space booms, deployable antennas, etc. Nevertheless, the inverse of the inertia matrix shall be calculated on the run, since it varies in time. This can be somewhat slow, mainly considering the high degree of freedom in dynamics of satellites with articulated panels and reaction wheels. In this case a more complex but fast set of equations in terms of the angular velocity is the best choice. This work describes the attitude dynamic equations expressed in angular velocities for satellites with appendages (solar panels, telescopes, directional antennas, etc.) and reaction wheels. They can be easily extended, however, to include nutation dampers and robotic arms. As usual, this formulation requires the knowledge of the acting torque in the connection joint, which is not always understood, modeled or known. So the dynamic model presented in this work uses the angular acceleration instead of the torque at the joint, which simplifies and reduces the order of the differential equations. Simulation results are presented, emphasizing the difference between the behavior of rigid body and articulated rigid bodies.

Reaction wheel is one of the actuators for spacecraft attitude control, which generates torque by changing an inertial rotor speed inside of the wheel. In order to generate required torque accurately and estimate an accurate angular momentum, wheel speed should be measured as close to the actual speed as possible. In this study, two conventional speed detection methods for high speed motor with digital tacho pulse (Elapsed-time method and Pulse-count method) and their resolutions are analyzed. For satellite attitude maneuvering and control, reaction wheel shall be operated in bi directional and low speed operation is sometimes needed for emergency case. Thus the bias error at low speed with constant acceleration (or deceleration) is also analysed. As a result, the speed detection error of elapsed-time method is largely influenced upon the high-speed clock frequency at high speed and largely effected on the number of tacho pulses used in elapsed time calculation at low speed, respectively.

In this paper, a passive fault tolerant attitude control is investigated for a nano-satellite with three magnetorquers and one reaction wheel in the presence of actuator constraints, external disturbances, and partial loss of actuator effectiveness as unknown actuator faults. For this purpose, a variable structure control approach is employed due to its ability to deal with the partial loss of actuator effectiveness and adjusting the maximum amplitude of the control signal by a control gain. A modified sliding variable is designed based on angular velocity and attitude errors with an adaptive parameter coefficient. The dynamics equation of the adaptive parameter is extracted during the stability proof of the closed-loop system. Changing the value of this parameter changes the effect intensity of the quaternion errors on the sliding variable and causes the attitude errors to maintain within a limited range. Applying this parameter to the modified sliding variable makes the controller perform well even if the adaptive parameter is significantly reduced. Furthermore, directly using the adaptive parameter in the structure of control law makes the controller generate a chattering-free torque. Comparing the simulation results with the results of two variable structure control methods designed in previous articles shows that its performance in attitude tracking commands has improved. In addition, Monte Carlo simulations with random actuator faults are performed to show the robustness of the proposed controller in the presence of actuator constraints and external disturbances.

Attitude control of combined service-target system in the post-capture phase has received great attention. A new attitude dynamics of the combined spacecraft with reaction wheels has been established in the author's former work, however, the measurement uncertainty in attitude and angular velocity and uncertainty in the reconfiguration matrix of reaction wheels have not been considered, which may cause huge impact on the system performance. In this paper, a novel combination of disturbance-observer-based dynamic surface control under measurement uncertainty and robust control allocation due to uncertain mass center is investigated for attitude stabilization of the combined spacecraft. Firstly, considering measurement uncertainty, inertia uncertainty, actuator fault and actuator saturation, a new attitude dynamics of combined spacecraft is established. Then, a virtual controller is designed and all the states in the closed-loop system converge to a small neighborhood of zero, where the lumped disturbance is compensated by two stable nonlinear disturbance observers and adverse effect of actuator saturation is addressed by a stable compensator. Finally, in consideration of uncertain location of mass center in the reconfiguration matrix, a LMI-based robust control allocation is employed to deal with the problem of distributing the three axis torques over the reaction wheels. Numerical simulations are presented to illustrate the effectiveness of the proposed method.

본 논문에서는 반작용 휠을 사용하는 위성체의 내고장 제어 문제를 다루고 있다. 고장은 구동기(반작용 휠)에만 발생한다고 가정하였으며, 고장에 대해서도 자세제어 성능을 유지할 수 있는 제어알고리듬을 제안한다. 자세제어 루프는 각속도 제어 루프를 포함하고 있으며, 각속도 제어 루프에 시간지연제어 기법을 적용하여 고장에 대응할 수 있도록 한다. 제안한 제어기의 안정성 조건을 제시하며, 시뮬레이션을 통해 다양한 형태의 고장에 대해 안정된 자세 추종 성능을 검증한다. This paper considers a fault tolerant control problem for a spacecraft using reaction wheels. Faults are assumed to be inherent to only actuators(reaction wheels) and a control algorithm to accommodate actuators' faults is proposed. An attitude control loop includes an angular velocity control loop. The time delay control method is used to make a spacecraft follow the command angular velocity and to accommodate actuators' faults. A stability condition for the proposed algorithm is derived and the performance is demonstrated by computer simulations.

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

A LTHOUGH the past three decades have witnessed several important developments in attitude stabilization control [1,2], there still remain certain open problems that are of great theoretical and practical interests, such as actuator constraint. Because of physical limitations, typical control actuators (e.g., reaction wheels and thrusters) for attitude control have an upper bound on the torque they can generate. If this issue is not solved during attitude maneuvering, then the attitude system will continue issuing its commands that may no longer be achievable by the satellite, and the required control will quickly saturate actuators. This will subsequently destabilize and even lead to loss of control of the satellite. Therefore, actuator constraint is one key issue that needs to be addressed. At present, a great deal of attention has been focused on the stability analysis and the attitude control design with saturated actuators [3–6]. Another problem that needs to be addressed is fault tolerant control for the attitude system. Any component or system fault can potentially cause a host of economic and safety problems. As a result, online and real-time fault tolerant control (FTC) design for satellites has received considerable attention [7]. Generally speaking, FTC can be classified into two types: passive and active [8]. The passive FTC is designed to be robust to a class of possible faults. It does not require fault detection and diagnosis (FDD) schemes. In an attempt to develop attitude controllers that can accommodate faults, extensive studies on the passive FTC control system for satellites have been performed [9–11]. The active FTC can react to fault events by using a reconfigurationmechanism. The active FTC relies on the availability of an FDD mechanism that gives, in real time, information about the nature and the intensity of the fault. The design of active FTC, especially the fault detection and diagnosis scheme for satellite attitude, has attracted considerable interest [12–14]. Although the methodologies including FDD and control law reconfiguration for the control-type faults of satellite attitude system have been studied extensively (as seen in the preceding references), the problem that is not solved yet is the effective and practical strategies for handling severe failure cases combining with actuators constraint. Lack of such effective FTC strategies may lead to substantial performance deterioration and even the loss of a satellite. The clear knowledge of the link between control reconfiguration and the actuator input constraint is still lacking. Because the control reconfiguration at different time instants with different types of faults has greatly different effects on control actuators saturation, there is no explicit and simple relationship between them, which makes this problem even more difficult. This leads to an important practical problem: how tomake use of the control reconfiguration on available control redundancy. Motivated by this problem, a novel attitude control approach is presented in this paper; an observer-based FDD scheme combining a sliding mode controller by using the estimated information is designed in a framework of active FTC system. The main contribution of this study, relative to otherworks, is that actuator faults, actuator constraints, and external disturbances are simultaneously addressed in an attitude FTC system design.

The Jupiter Magnetospheric Explorer (JMEX) is a UV observatory operating in an earth orbit proposed as part of NASA's Small Explorer (SMEX) class of missions. To meet mission requirements the residual jitter portion of the imaging error budget is set at 0.079 arcsec (3σ) over a 33.3 ms frame integration time and 0.01 arcsec (3σ) for all frequency content higher than 15 Hz. These requirements are challenging for a small, low cost mission and require some innovative system solutions to achieve these goals. The solution, discussed in the paper, was to combine several jitter rejection techniques fine-balanced reaction wheel mounted on an isolation assembly, post processing using science images and reaction wheel momentum control. This paper focuses primarily on meeting the high frequency portion of the requirements. To facilitate system performance verification, we leveraged an integrated model toolset, EOSyM (End-to-end Optical System Model), developed and used on various other advanced space-based missions over the last 9 years. Starting with individual subsystem models for the reaction wheel disturbances, the coupled payload/ spacecraft structural dynamics model, and the optical design, we were able to evaluate the end-to-end LOS performance under varying reaction wheel speeds. At the end we found that the requirements could be met by maintaining the reaction wheels operating range within a well-defined speed band. This paper describes the mission, the technical challenges, the integrated model, and system performance results.

Flutter analysis of a rigid-flexible coupled composite space structure with momentum wheels under thermal load

Attitude jitter is a common phenomenon in the application of high resolution satellites, which may result in large errors of geo-positioning and mapping accuracy. Therefore, it is critical to detect and compensate attitude jitter to explore the full geometric potential of high resolution satellites. In this paper, a framework of jitter detection and compensation for high resolution satellites is proposed and some preliminary investigation is performed. Three methods for jitter detection are presented as follows. (1) The first one is based on multispectral images using parallax between two different bands in the image; (2) The second is based on stereo images using rational polynomial coefficients (RPCs); (3) The third is based on panchromatic images employing orthorectification processing. Based on the calculated parallax maps, the frequency and amplitude of the detected jitter are obtained. Subsequently, two approaches for jitter compensation are conducted. (1) The first one is to conduct the compensation on image, which uses the derived parallax observations for resampling; (2) The second is to conduct the compensation on attitude data, which treats the influence of jitter on attitude as correction of charge-coupled device (CCD) viewing angles. Experiments with images from several satellites, such as ASTER (Advanced Spaceborne Thermal Emission and Reflection Radiaometer), LRO (Lunar Reconnaissance Orbiter) and ZY-3 (ZiYuan-3) demonstrate the promising performance and feasibility of the proposed framework.

The Attitude and Orbit Control System of the 3 axis stabilised spacecraft (S/C) XMMNewton was designed to use three reaction wheels (RW) with a fourth wheel in cold redundancy. The 3 wheel configuration has been used for more than 13 years for the momentum management, to re-orientate the S/C with slews in order to achieve the required scientific pointing. With 3 wheels, frequent reaction wheel unloading is used both to compensate the external torque and to re-orientate the angular momentum to allow the slew execution, which results in fuel consumption of around 6 Kg per year. A fourth wheel in control gives the advantage of a higher momentum envelope and introduces the null space in the wheels matrix, allowing a wheel speed variation without changing the total angular momentum. With four wheels the reaction wheel bias will compensate only for the accumulation of angular momentum due to the external torques, resulting in saving a significant amount of fuel and extending the mission lifetime. The use of four reaction wheels implied the development of new attitude control algorithms and a software change on board. A design phase of this change was successfully completed and the project went through the implementation