Improvement and verification of satellite dynamics simulator based on flight data analysis

T HIS Note proposes a new strategy for the design of the attitude determination and control system (ADCS) of reconfigurable spacecraft, with a specific application example to so-called nanosatellites. These spacecraft adhere to the CubeSat standard [1], and the natural discretization of the overall volume into cube units makes them a popular platform for research into spacecraft reconfiguration [2]. The proposed design strategy, called location-scheduled control (LSC), seeks to reduce the development time required for the ADCS of a reconfigurable spacecraft conducting scientific research. In traditional ADCS design, mission science requirements levy pointing requirements onto the ADCS, which drive sensor and actuator selection. Once the ADCS hardware has been chosen, an appropriate control algorithm is selected and tested. In contrast, the location-scheduled approach to ADCS design is a two-phase process. In the first phase, a general purpose self-contained ADCS hardware unit is designed by selecting a suite of sensors and actuators that are appropriate for a class of science missions. Controller gains and ADCS location are then optimized in simulation for a number of generic science payloads, such as those depicted in Fig. 1, in which payload cube (P) layouts dictate possible locations for the ADCS (cubes A or B). In the second phase of the process, one of the previously tested configurations is selected based on the needs of the current science mission. Some amount offine gain tuningwill inevitably be required, but if a member of the catalog of spacecraft configurations satisfies the needs of the science payload, the developer need not return to the first phase of the design process. Although the LSC approach is applicable for many reconfigurable spacecraft, thiswork considers a nanosatellite ADCSdesign problem to demonstrate this methodology. The objective of the design problem is to find the optimal combination of ADCS unit location and control law gainvalues for a nanosatellite performing a detumble maneuver in which the spacecraft has high initial angular velocity and arbitrary initial orientation. Using the CubeSat standard, a 10 10 10 cm cube is designated as one unit of volume (abbreviated 1U), and complete spacecraft are created from these cubes. Traditional CubeSats are typically no larger than 3U in volume, and they have largely been passively controlled via permanent magnets [3,4], gravity gradient booms, or aerodynamic stabilization. This Note focuses on spacecraft that are 6U in volume (as in Fig. 1), with overall dimensions of 10 20 30 cm. This form factor creates greater volume for both complex science payloads and attitude control hardware and is currently under development at NASAAmes Research Center. A nonlinear control law introduced in [5] is used to control the nanosatellite using actuators for which the control authority is restricted by a known upper bound.

Attitude determination & control system design for gravity recovery missions like GRACE

Flight simulators can recreate aircraft flights for flight training and aircraft design. This paper reviews existing engineering flight simulators and introduces the development of a multi-purpose reconfigurable engineering flight simulator at Ryerson University. The multi-purpose engineering flight simulator named the Ryerson Fixed Base Simulator (RFBS) has been designed and built to teach and initiate research projects in the area of aircraft design, flight simulation, pilot training, and flight data analysis. It consists of three 46 inch high definition screens and six 22 inch touch screen panels to represent the instrument panel, the centre console, and the overhead panels of an actual aircraft flight deck. Several low-cost, commercial flight simulation software were tested and X-Plane was selected as the main flight simulation tool. This paper also introduces a research that utilizes the RFBS and a commercial flight analysis software. The objective of the research was to develop a flight data conversion methodology. A case study was commenced to verify the work flow of flight data generation and analysis with an example of hard-landing analysis.

Development and experimental validation of a multi-algorithmic hybrid attitude determination and control system for a small satellite

Fault detection and isolation algorithms for LEO satellite attitude determination and control system using an approach for checking the statistical characteristics of EKF innovation sequence are proposed. The fault detection algorithm is based on statistic for the mathematical expectation of the spectral norm of the normalized innovation matrix of the EKF. This approach permits simultaneous real-time checking of the mathematical expectation and the variance of the innovation sequence and does not require a priori information about the faults and statistical characteristics of the system in fault cases. In this study an attitude estimation and control system for LEO satellite is proposed. To determine the attitude of the satellite, this system use algebraic method (two-vector algorithm). As a reference direction, the unit vectors toward the Sun, the Earth’s center, and the Earth magnetic field are used. Thus, it includes three different two-vector algorithms based on using the Earth’s magnetic field – the Sun vector, the Earth’s magnetic field – nadir vector, and nadir vector – the Sun vector couples. In order to increase the attitude determination accuracy, the redundant data processing algorithm, based on the Maximum Likelihood Method, is used. An extended Kalman filter has been developed for nonlinear rotational dynamics estimation of LEO satellite. Failures in the sensors affect the characteristics of the innovation sequence of the EKF. The failures that affect the mean and variance of the innovation sequence have been considered. The application of the proposed fault detection algorithm to the LEO satellite attitude determination and control system has shown that, sensor fault detection by the presented algorithm is possible in real time. Assuming that the effect of the faulty sensor on its channel is more significant than on the other channels, a sensor isolation method is presented by transforming n-dimensional innovation process to n one-dimensional processes. The simulations, carried out on a nonlinear dynamic model of the rotational motion of LEO satellite, confirm the theoretical results. The future work is to investigate the faults affecting the attitude dynamics e.g., actuator faults and to perform the integrated sensor/actuator FDI via innovation approach.

The Attitude Determination and Control System of the Picosatellite UWE-3

Design and on-orbit performance of the attitude determination and control system for the ZDPS-1A pico-satellite

The redundancy allocation problem is to find an optimal allocation of redundant components by considering the set of resources and cost constraints. In this research, the satellite attitude determination and control system is studied and its components is introduced, then the reliability of this system is modeled and optimized based on a mathematical approach based on the redundancy allocation. The model studied in this paper is about the structure of a series-parallel system that is in the exact mode of a satellite attitude determination and control system. In this paper, a new approach for modeling and optimization is presented. The mathematical model presented in this paper is into the class of mixed integer non-linear programming (MINLP). Solving these problems is very important for various researchers due to the high mathematical complexity. In this research, a heuristic method is used to the problem exact solution, the results of which have been reported.

Assessing attitude error of FORMOSAT-3/COSMIC satellites and its impact on orbit determination

Optical navigation on a CubeSat must rely on the best extraction of the directions of some beacons from on-board images. We present an experiment on OPS-SAT, a CubeSat of the European Space Agency (ESA), that will characterize an on-board algorithm to this aim, named Angle-based Correlation (AbC). OPS-SAT is a 3-unit CubeSat with an Attitude Deter-mination and Control System (ADCS) and an imager that have proved their reliability with typical performance at CubeSat scale. We selected a few star-fields that all present enough visible stars within a 10º field of view. When our experiment is run, OPS-SAT is pointed to the most convenient star-field at that time. There, the star-field is imaged and subwindows are extracted from the image around the expected location of each star, based on the attitude-quaternion reported by the ADCS. The AbC reconstructs the absolute direction of the central body, in principle unknown, which is the pointed known star in the experiment. The method intensively uses the quaternion algebra. The beacon location is first consolidated in the field of view with the AbC. Then, the field of view is finely positioned against the sky, again with the AbC. A covariance is associated with the found beacon direction. Our experiment with OPS-SAT manages the pointing and the imager, and processes the taken images. Then, it downloads the on-board computed absolute directions and their covariances, to be compared with the actual directions. After a campaign of intensive use of the experiment, the statistical performance of the algorithm will be established and compared to the on-board computed covariances. As a bonus, an assessment of OPS-SAT's inertial pointing stability will be available. The AbC can theoretically get rid of the Attitude Control Error (ACE) of the platform and of the Attitude Knowledge Error (AKE) estimated by the ADCS, and potentially converge to sub-arcsec measurements, whereas 15 to 90 arcsec for AKE are expected on CubeSats and imager pixel size ranges in 10 to 15 arcsec. Results will be available from 2022, thus, instead, we illustrate here possible applications for commissioning and autonomous operations.

Spacecraft attitude control systems rely on reaction wheels as the primary means of precise three-axis attitude control. Faults in these reaction wheels might lead to system instability and, in severe cases, mission failure. This paper presents advanced machine learning-based techniques for the detection, isolation, and identification of reaction wheel faults in spacecraft.The proposed approach leverages advanced data analytics and machine learning algorithms to analyze sensor data from the reaction wheels, enabling early detection of faults and effective isolation of the faulty component and identify the types of faults detected, specifically, voltage, current and temperature faults.Three-axis controlled satellite high-fidelity models are simulated to generate data for both nominal and faulty states of RW. The simulated data is employed with the FDII approach. The generated data is passed into five different machine learning classifiers, the isolation and identification results are verified via cross-validation. The proposed techniques is tested on three defined datasets using the three-orthogonal RW configuration to verify its robustness. The results show that the system has higher isolation and identification accuracy when compared to other studies that used various methodologies.

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.

The Cyclone Global Navigation Satellite System (CYGNSS), which was recently selected as the Earth Venture-2 investigation by NASA's Earth Science System Pathfinder (ESSP) Program, measures the ocean surface wind field with unprecedented temporal resolution and spatial coverage, under all precipitating conditions, and over the full dynamic range of wind speeds experienced in a tropical cyclone (TC). The CYGNSS flight segment consists of 8 microsatellite-class observatories, which represent SwRI's first spacecraft bus design, installed on a Deployment Module for launch. They are identical in design but provide their own individual contribution to the CYGNSS science data set. Subsystems include the Attitude Determination and Control System (ADCS), the Communication and Data Subsystem (CDS), the Electrical Power Supply (EPS), and the Structure, Mechanisms, and Thermal Subsystem (SMT). This paper will present an overview of the mission and the avionics, including the ADCS, CDS, and EPS, in detail. Specifically, we will detail how off-the-shelf components can be utilized to do ADCS and will highlight how SwRI's existing avionics solutions will be adapted to meet the requirements and cost constraints of microsat applications. Avionics electronics provided by SwRI include a command and data handling computer, a transceiver radio, a low voltage power supply (LVPS), and a peak power tracker (PPT).

Attitude control system plays the important role for to maintain the satellite to desired orientation. To control the satellite it is necessary to do the attitude stabilization. Attitude stabilization achieved by Star sensor, sun sensor, Earth sensors. Attitude control is mainly use for antenna pointing accuracy, camera focus to earth surface and solar panel pointing toward sun. Due to tumbling effects, satellite will rotate all the direction in the space. To maintain the orientation of the satellite it is necessary to design the attitude determination and control. Satellite consider as the rigid body. Inertia matrix describes the rigid body dynamics. The orientation of the satellite determine by Euler angle and Quaternion. The Low earth orbiting satellite will have enormous amount of aerodynamic drag stinking the satellite body and gravitational attraction another problem. Because of that satellite dwell, time is reduced. It means satellite more time spending particular part of the earth. The attitude estimation is measures by the orientation of vectors. Attitude estimation means to find the position and orientation of flying object with respect to the fixed reference of reference. Vector remains considered in the frame of reference to compute for find the orientation of the body of the satellite in the inertial reference system. The Earth is an inertial reference frame, and Satellite is a body frame. Attitude sensor used to measure the satellite orientation in the reference frame. This will help in accurately predicting the orbit deviation and a control system to correct if any by providing the satellite momentum means ‘mass in motion’ changes in a body rapidly in Low earth orbit due to centripetal force acting on a satellite. Attitude control system (ACS) need the numerical simulation to find the required torque demand by the help of difference between reference input (Attitude) signal and feedback signal measure by attitude sensor to trim the control surface maintain the actuator required orientation . The results will consist of two parts the first part consisting of the attitude estimation using Euler angle and Quaternion method, second part consist of estimate the control torque from magnetic torquer and error estimation using non-linear filter (Unscented Kalman Filter) with MATLAB simulation. Â

Flight data can not only be applied to safety analysis, condition monitoring and model establishment, but also play a very important role in improving the effect of theory teaching and students' own ability. Based on flight simulation software and visual programming language, a flight data application system is developed in this paper. The system includes three modules: flight status monitoring, verification experiment and data customization. It can realize the objective evaluation of flight skills, the analysis and verification of flight data of theoretical knowledge, as well as the functions of classroom teaching demonstration and independent design experiment of teachers and students. It solves the current situation of insufficient application of simulated flight data in theoretical teaching and students' ability training. The teaching practice shows that the system can improve students' flight technology, deepen their understanding of abstract knowledge, stimulate their independent exploration ability, and assist teachers' classroom teaching, which has high application value.