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

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.

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.

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.

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.

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.

This paper advocates a new approach to satellite software design based on object-oriented framework technology and describes early results from a project for the European Space Agency (ESA) to design a software framework for satellite attitude and orbit control systems (AOCS). Frameworks are collections of components with pre-defined cooperations among them. They make architecture (as opposed to mere code) reuse possible. The framework concept is being tested in a redesign of the AOCS software. This paper illustrates it by describing the implementation of telecommand handling, telemetry handling, and operational mode management.

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.

LEO MYRIADE AOCS ADAPTION TO GTO FOR SPIRALE MISSION

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

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.

A feasibility study is presented for using micro Pulsed Plasma Thrusters (μPPT) on a 3U CubeSat to perform attitude control to meet pointing requirements at altitudes between 400 and 1000 km. The orbits and attitude control are consistent with science missions for ionospheric and solar observations. The positioning of the thrusters is such that they provide full, three-axis rotational control while maximizing the efficiency of each pulse. The aerodynamic, magnetic, solar radiation, and gravity gradient disturbance torques are included in the formulation. Sensor modeling is also included. The study involves realistic power constraints anticipated on the 3U CubeSat. The paper presents the overall design of the 3U CubeSat and focuses on the attitude determination and control (ADC) approach and implementation. The paper also introduces the control strategy and numerical simulation results for stabilization, pointing and spinning applications.

The Clementine spacecraft, launched in early 1994, recently completed a two-month lunar mapping mission in which it recorded high-resolution images over the entire surface of the moon. The attitude determination and control subsystems developed for the mapping phase of the mission are described. The vehicle's attitude determination subsystem consisted of a simple fixed-gain Kalman filter, uncoupled about each spacecraft axis, combining rate inputs from two solid-state optical gyros and attitude inputs from two wide field-of-view star trackers. The attitude control subsystem consisted of a simple proportional-integral-derivative algorithm operating on quaternion and rate feedback signals to provide pointing, tracking, and slewing functions. The primary actuating devices were three reaction wheels used for feedback error compensation and 12 thrusters used for occasional momentum dumping. During lunar mapping the attitude determination and control subsystems maintained vehicle pointing to within a three-axis rss value of 0.05 deg. On-orbit results are provided from outputs of the attitude determination filter, the lunar nadir-tracking controller, and the large-angle slew generator. Additionally, results are provided from a novel inertia estimator, which utilized sensor outputs during two slew maneuvers.

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.

In designing the attitude determination and control system (ADCS) of a satellite, the system cannot be modeled as a dimensionless point object. A practical object has physical size and thus rotates as well as translates. Such bodies are termed rigid bodies. Its dynamics is characterized by six degrees of freedom. Three degrees of freedom represent the motion of the center of mass. The other three degrees address the rotational motion of the object. The 6-DoF must be dynamically controlled. Traditionally, attitude determination involved using both rate gyros and star trackers. Rate gyros are prone to mechanical failure. Hence it is proposed to study the feasibility of replacing rate gyros with star sensors for attitude determination. A Kalman filter provides an optimal means to estimate the state of the system. Results show that the rates can be estimated to within 10/sup -7/ rad/s which is as good as from a gyroscope.

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