A component framework for satellite on-board software

In this paper, an innovative fault-tolerant distributed design of a satellite Attitude and Orbit Control System (AOCS) is proposed. First, the system functionality and application tasks are specified, and measurable performance parameters are defined. Then a novel multiprocessor system-onchip (MPSoC) AOCS computer is designed, which is configured to operate as a fault-tolerant distributed computing system. Next the distributed MPSoC AOCS is prototyped using XILINX Zynq FPGAs. The prototype is fully tested as hardware in the loop with a MATLAB AOCS modeland a fault injection system. The experimental results confirm that the fault-tolerant distributed computing scheme outperforms existing on-board approaches in terms of resilience to processor failures and computational performance.

In this study, an optimisation-based approach is proposed for the robustness analysis of an attitude and orbit control system (AOCS) for flexible satellites. Several optimisation methods, including local gradient-based algorithms, global evolutionary algorithms and hybrid local/global algorithms are applied to the problem of analysing the robustness of a full-authority multivariable controller with respect to several frequency and time domain performance criteria, for a 6 degree of freedom simulation model of a satellite with large sun shields. The results of our study reveal the advantages of optimisation-based worst-case analysis over traditional Monte Carlo simulations for systems with flexible dynamics. In particular, it is shown that frequency domain analysis can provide useful guidance for the formulation of subsequent time domain tests, and that hybrid local/global optimisation algorithms can produce more reliable estimates of worst-case performance, while also reducing the associated computational overheads. The proposed approach appears to have significant potential for improving the industrial flight clearance process for next-generation high-performance satellite control systems.

The development of space power engineering is one of the well-known lines in rocket and space science and innovative technologies which attracts the attention of many scientists and researchers. Engineering solutions in space-based solar power plant design and wireless space-to-Earth and satellite-to-satellite power transmission and power spacecraft control methods have been substantiated theoretically to sufficient depth. However, despite this, there is a need to improve methods for and approaches to the development of an optimal design methodology for power spacecraft. A way to improve existing approaches to the development of space-based solar power plants and power satellites may be the use of mobile control methods in the development of an attitude and orbit control system. Such methods allow one to reduce power consumption for control operations. The goal of this paper is to study the features of mobile control and construct a methodology for the development of solar power satellites’ attitude and orbit control system (AOCS) using mobile control algorithms. The paper considers the features of mobile control algorithm synthesis for the attitude control and stabilization of solar power spacecraft (solar power plants and power satellites). Power spacecraft control tasks are classified, and the expediency of using mobile control methods is justified. An analysis is made for the stability problem that arises in controlling power spacecraft with flexible elements. The paper presents methodological recommendations on determining the AOCS design parameters for space-based solar power plants and power spacecraft for wireless satellite-to-satellite power transmission. This methodology may be used in power satellite development.

The FORMOSAT-5 (FS-5), the fifth space program initiated in 2008, is currently developed by National Space Organization (NSPO) of the National Applied Research Laboratories (NARL) in Taiwan, Republic of China (ROC). The satellite will be deployed into a sun-synchronous orbit with 720-km altitude and 98.28o inclination angle. It is a threeaxis attitude stabilized spacecraft providing high precision attitude determination and control for a high resolution Remote Sensing Instrument (RSI). To meet the 0.012 deg, 3sigma, per axis, absolute bus attitude knowledge requirement flown-down from RSI for meeting its geo-location accuracy, the micro-Advanced Stellar Compass (ASC) consisting of three Camera Head Units (CHUs) along with four single-axis fiber-optics gyros is mounted on a stable optical bench common to the RSI structure. The ASC is a highly advanced and fully autonomous star tracker produced by the Measurement & Instrument System (MIS) Section of the Orsted Department of the Technical University of Denmark (DTU). The gyroscopes considered in this mission are the fiber-optics rate sensors, -FORS6U, manufactured by Litef, a Northrop Grumman branch in Germany. The -FORS6U was originally designed to meet the requirements of a wide range of air, land, and sea applications, and was later certified by NSPO for space applications after a series of environmental tests. The primary spacecraft attitude and rate determination for the spacecraft Attitude and Orbit Control System (AOCS) will be provided by a gyro-stellar attitude determination system, which utilizes attitude data provided by three CHUs and rate data provided by the four single-axis fiber-optics gyros. To account for the possibility of more than one gyro failures due to lack of space-flight heritage, a gyro-less attitude and rate determination algorithm is developed as a back-up for the AOCS design. This paper describes the detailed development of a gyro-less attitude and rate determination algorithm which uses the computed gyro data (numerical gyro data) instead of physical gyro data as in a typical gyro-stellar attitude determination system. The numerical gyro data are computed using the assumed spacecraft dynamics (or Euler equations of motion). By perturbing the spacecraft kinematic equations, the linearized attitude error equations are obtained. Similarly, by perturbing the spacecraft dynamic equations, the linearized rate error equations are obtained. With the obtained linearized attitude error equations and the linearized rate error equations, a reduced-order (6-state) Extended Kalman Filter (EKF) providing spacecraft attitude and rate estimates is then implemented in the algorithm. The developed gyro-less attitude and rate determination algorithm was incorporated and tested in a 6-DoF nonlinear, high-fidelity simulation model developed for the FS-5 program to assess its performance. Its performance estimates and sensitivities to spacecraft inertia tensor uncertainties, reaction wheel momentum uncertainties, and torque uncertainties produced by torque-rods during normal operations will be presented in the paper.

The magnetic field data from the GOES satellite must be corrected on the ground for the contribution from the roll and yaw torquer coils which comprise important elements of the spacecraft attitude and orbit control system (AOCS). The correction may be calculated from measurements of the torquer current which are telemetered at the same rate as the magnetic field data or from torquer input command data. The baseline plan was to allow torque current changes only at infrequent intervals of once per 20 minutes since it was suspected that torquer switching noise might be a problem. However, for improved navigation and registration performance in the radiometric instruments, the actual mode of operation has been to calculate the torquer current drive more frequently and to rate limit the current changes to one commanded increment per step. The magnetic signature of the 3.2 ma step is an much as 1.2 nanoTeslas. Both the current level and the telemetered value change essentially instantaneous when compared to the frequency response of the magnetometer, which incorporates a 0.5 Hz bandwidth five- pole Butterworth anti-aliasing filter in each data channel. Straightforward correction of the data using the current telemetry therefore results in objectionable noise spikes in the derived bandwidth-limited output data. The problem is complicated by the fact that the clock which governs the AOCS current commands to the torquers is not synchronized with the telemetry unit clock which governs magnetometer data sampling. Fortunately, it has been possible to incorporate data in the spacecraft telemetry which provides improved information for determining the time of a current step in the telemetry unit time base as well as the torquer input command data mentioned as one method of compensation. This paper describes the development of a numerical filter used in ground processing of the telemetered current to simulate the five-pole Butterworth response of the magnetometer output filter. Since the timing information is determined to an accuracy of 1 part in 8, one of eight sets of filter weights is selected based on the telemetered timing information. Examples are given of the improvement in output data quality resulting from the technique.© (1996) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Visual servoing for coordinated precise attitude control in the TOM small satellite formation

The spacecraft controllability of the angular motion is possible only with operability of the attitude and orbit control system (AOCS) of the spacecraft, sensors, actuators and the spacecraft power system. However, there is a rather significant probability of failure of this equipment during the operation of the spacecraft. This is especially observed after half of the spacecraft's lifetime or because of emergency situations. There is a problem which is connected with providing the maximum performance of the AOCS in case of partial failures of their actuators (reaction wheels (RW), magnetorquer rods (MGTR), etc.). Thus, the purpose of this work is the development and synthesis of special algorithms for spacecraft angular motion control in the emergency situations which are connected with RWs partial failures and restrictions of onboard electricity consumption. The approach of synthesis of this control algorithms is based on using mobile control methods which allow to reserve RWs by MGTRs. There are different variants of control loops depending on MGTRs turning on combinations. There were proposed two types of control switching functions: time-periodic and switching by deviation. Also was proposed a methodology of controller synthesis using these switching functions. Using this methodology and computer simulation, it was shown the possibility of providing angular nadir orientation and stabilization of the spacecraft with maximum 1−1.5 deg error in case of time-periodic switching functions implementation. Switching by deviation allows to reduce onboard electricity consumption for 25−30 % comparing with using time-periodic switching. However, the accuracy of stabilization significantly lower in case of switching by deviation. Considering these estimates, the corresponding methodological recommendations were formulated for use switching functions depending on emergency

Spacecraft play an increasingly important role in various areas of modern society, such as telecommunication, earth observation, and space exploration. It is estimated that there have been more than 7000 spacecrafts launched all over the world. Despite rigorous testing many of these spacecraft fail on orbit due to various reasons [1], which consequently often lead to the failure of the whole mission. According to [2], over 30% of spacecraft failures occur at the subsystem level of the Attitude and Orbit Control System (AOCS). Moreover about 50% of the AOCS failures are attributed to actuator errors. The purpose of this paper is to present an actuator fault-tolerant attitude control.

LEO MYRIADE AOCS ADAPTION TO GTO FOR SPIRALE MISSION

Due to increasing commercialisation in satellite business modern companies have to create proposals quicker and cheaper than ever before. To cope with those demands a model based design approach is currently implemented at the satellite division of DaimlerChrysler Aerospace.The attitude and orbit control system (AOCS) is one of the key subsystems within a satellite, as it's heavily interfacing to all other subsystems.A tool was developed which supports the AOCS‐designer in selecting suitable COTS‐components based on functional requirements. For each functional requirement (product‐function) a set of reference solutions is identified. Based on the designers selection of those reference solutions, a pre‐selection of suitable COTS‐components is being filtered from a component database.After the designer has selected the final set of components, the tool provides a compliancy check to the initial requirements.

On-orbit applications, such as active debris removal, satellite refueling, maintenance and satellite health diagnosis require the ability for low-cost spacecraft to closely inspect other orbital objects in a non-cooperative manner. During this kind of mission, the relative navigation process becomes of critical importance for guaranteeing safe and collision-free proximity operations and maneuvers. The selection of relative navigation sensors appears being a critical task as it might drive the design choices in other subsystems (e.g., Attitude and orbit control system (AOCS), payload, power supply). This paper aims at studying the application of a Multidisciplinary Design Optimization (MDO) method to the design of AOCS and Navigation subsystems under small satellite constraints. The porposed MDO process is based on the optimization of navigation performance and mass reduction while respecting volume and power constraints for orbital close rendezvous. The navigation chain, including sensor simulation and navigation filter is simulated and integrated into the design cost function. The proposed MDO process based on a Genetic Algorithm (GA) and on an Extended Kalman Filter (EKF) aims at simplifying satellite design by determining a set of optimal admissible sensor combinations despite contradictory objectives on navigation and payload accuracy, mass reduction, power consumption and volume. Possible key advantages of the inclusion of the relative navigation subsystem within MDO process are the reduction of the design process time, the automation and optimization of the navigation architecture while respecting volume and power constraints of small satellites. A demonstration of the effectiveness of the proposed MDO method is provided on the benchmark of AVANTI mission.

The recent decades have witnessed active studies in fault detection and isolation (FDD) of the small satellites subject to actuator failures and attitude coordination control of spacecraft formation flying (SFF) which plays a significant role in missions such as synthetic-aperture radar imaging. However, because of the time delay in communication links, the performance of coordination control is not always satisfactory. In this paper we propose a method to estimate the actuator fault as well as the input time delay at the same time. First of all, the model of the altitude control system of the small satellite with actuator fault and the model of the altitude control system with actuator fault and the input time delay are presented respectively. Based on the Two-Stage Kalman Filter algorithm, the fault and input time delay of the altitude control system of the small satellite can be estimated. At last, the simulation of the algorithm in a satellite attitude control system with actuator fault and input time delay will illustrate the effectiveness of the proposed approach.

In-flight Assessment of Spot Attitude and Orbit Control

Relying on the organic organization and cooperation of the multifunctional organs in the body, the organism completes the basic form of life. Attitude Orbit Control System of Satellite (AOCS) is composed of multiple sensors and actuators. It is effectively organized and processed by the on-board computer to complete the attitude and orbit control tasks of the satellite platform. This article compares AOCS to “organisms” that require long-term health diagnosis and treatment and organically decomposes their nominal functional blocks. Then an AOCS quantitative health assessment method is proposed that the health quantitative diagnosis system designed based on. And it can calculate the satellite AOCS quantitative health status score to support the maintenance of the on-orbit satellite AOCS efficiently.

Space Qualification is a vital component of successful spacecraft design. The entire scope of the space qualification process is extremely large. This thesis will focus primarily on the Attitude Control System and Space Environment model, with the intent of furthering the verification of the design for the Joint Australian Engineering Satellite (JAESAT). The Attitude Control System proposed for use aboard JAESAT is a combination of permanent magnets with hysteresis rod damping. This system has proved effective on many previous satellite missions, however analysis must be undertaken to optimise the control provided for the individual design proposed. The Space Environment will affect the operation of many of the subsystems aboard JAESAT. Knowledge of the magnitude of such affects is essential to prevent complications through improved design. To do this requires an accurate model of the Space Environment local to JAESAT. Several models for different environmental characteristics are presented as part of this dissertation. Further detailed modelling was performed using the Space Environment Information System, an online software package provided by the European Space Agency. Some of the many results that have been obtained from the Space Environment Information System are presented in graphical format. A detailed model of the Attitude Control System was constructed using Matlab. The environmental disturbance torques were calculated using the environmental conditions previously modelled. Several simulations were conducted yielding mixed results. Generally, it was found that a permanent magnet strength of 14 A.m2 and spin rate of 0.4 revolutions per minute is sufficient to control JAESAT to within 1° of the magnetic field vector. Numerical simulation of the hysteresis rod damping however, has been shown to be ineffective and additional testing is required to successfully implement the hysteresis effects in the system simulation. An analysis of the model revealed that there were minimal sources of error present in the simulation process. The only significant source of error was due to the numerical integration of the equations associated with the satellite motion. Another possible source of model inaccuracy was due to the variation between the numerically simulated space environment and the actual space environment. This dissimilarity however, will only result in small changes of satellite orientation and will not have any effect on the satellite attitude. In the future, the Space Environment and Attitude Control System model that was constructed should be updated to reflect any changes in satellite design and orbit parameters. Simulations of the final Attitude Control System hardware should also be conducted in order to show that adequate attitude control is maintained throughout a period of several orbits. The hysteresis rods should be tested in order to determine their damping properties, and can then be incorporated in the satellite attitude simulations. Based upon numerous simulation results and research, it was concluded that the Attitude Control System that has been proposed was sufficient to provide the required level of attitude control for JAESAT. The successful verification of the Attitude Control System operation implies that the system is Space Qualified. Exact specifications for the Attitude Control System hardware are dependant upon the satellite orbit and design requirements. The model of the space environment is also sufficiently accurate to use in the verification of other satellite systems.