Attitude Determination and Control Systems

The importance of accurately pointing spacecraft to our daily lives is pervasive, yet somehow escapes the notice of most people. In this section, we will summarize the processes and technologies used in designing and operating spacecraft pointing (i.e. attitude) systems

An Attitude Control System for a Low-Cost Earth Observation Satellite with Orbit Maintenance Capability

UoSAT-12 is a low-cost minisatellite built by Surrey Satellite Technology Ltd. (SSTL), it is amongst other objectives also a technology demonstrator for high performance attitude control and orbit maintenance on a future constellation of earth observation satellites. The satellite uses a 3-axis reaction wheel configuration and a cold gas propulsion system to enable precise and fast control of its attitude, for example, during orbit manoeuvres. Magnetorquer coils assist the wheels mainly for momentum dumping. This paper describes the various attitude control modes required to support: 1) the initial attitude acquisition phase, 2) a high resolution imager payload during pointing and tracking of targets, 3) the propulsion system during orbit manoeuvres. The specific attitude controllers and estimators used during these control modes are explained. Various simulation and in-orbit test results are presented to evaluate the performance and design objectives. To improve the control and estimation accuracy, on-board calibration and alignment procedures for the sensors and actuators are utilised. Some calibration results and the resulting improvement in accuracy from these procedures are shown

Attitude Determination and Control Subsystem Design for a CubeSat

This project continues the design and testing of the Attitude Determination and Control Subsystem (ADCS) for a nano-satellite. The primary mission objective is solar X-ray spectroscopy using the Sphinx-NG instrument, which requires that the CubeSat fly in a high-altitude, polar, sun- synchronous orbit pointing to the sun with 1-2 degrees of accuracy. The ADCS requires gyroscopes, sun sensors, and a magnetometer for attitude determination. Attitude control is executed using magnetorquers as actuators. This project focused on the analysis of attitude determination algorithms and control policies to select the most efficient and accurate methods. After method selection, simulations of the ADCS were conducted, and research was performed concerning hardware testing for the ADCS

Three-axes attitude determination and control system based on magnetorquers for small satellites

In this TFM the attitude determination and control system for the 3Cat-2 nanosatellite will be implemented, including the sun sensors, 3 axes magnetometers and 3 axes gyroscopes to determine the satellite's attitude, and three axes magnetorquers as actuators. Algorithms will be tested in a Matlab-based simulation environment and then converted into C, to be executed in the Nanomind (on board computer) of 3Cat-2. Algorithm performance should be evaluated in an air bearing in the center of a 3 axes Helmholtz coils.This projects takes the desing and implementation of an attitude determination and control subsystem for the 3Cat-2 Mission. The system has to be able to correct the different perturbations and point the satellite to the desired orientation according to the working mode of the satellite. The subsystem has to be able to correct the different sensing perturbations and compute all the different attitude parameters to control the satellite. The design is divided in three different modes: Detumbling, Sun-safe, and Nominal. The Detumbling Mode is in charge of stabilizing the satellite rotation after the launch or when some other of the controllers has had a problem, and it induces a high rotation over the satellite. The Sun-safe is in charge of pointing the largest solar panels to the Sun to charge the batteries. Finally, the Nominal has to point the payload antennas to the Nadir (Earth Center) for performing the mission of the satellite. Each mode will make use of different sensors to save as much energy as possible. The work will be based on the usual algorithms used in different satellite mission and looking for a new and the first complex implementation of an Attitude subsystem for the Nanosatellite Laboratory

ALGORITHMS AND OPTIMAL CONTROL FOR SPACECRAFT MAGNETIC ATTITUDE MANEUVERS

This study focused on providing applicable control solutions for spacecraft magnetic attitude control system. Basically, two main lines are pursued; first, developing detumbling control laws and second, an improvement in the three-axis attitude control schemes by extending magnetic rods activation time. Spacecraft, after separation from the launching mechanism, experiences a tumbling phase due to an undesired angular momentum. In this study, we present a new efficient variant of the B-dot detumbling law by introducing a substitute of the spacecraft angular velocity, based on the ambient magnetic field data. This B-dot law preserves the orthogonality, among the applied torque, dipole moment and magnetic field vectors. Most of the existing variants of the B-dot law in the literature don\u27t preserve this orthogonality. Furthermore, the problem of minimum-time spacecraft magnetic detumbling is revisited within the context of optimal control theory. Two formulations are presented; the first one assumes the availability of the angular velocity measurements for feedback. The second formulation assumes the availability of only the ambient magnetic field measurements in the feedback; the latter is considered another optimal-based B-dot law. A reduction in detumbling time is fulfilled by the proposed laws along with less power consumption for the proposed B-dot laws. In magnetic attitude maneuvers, magnetic rods and magnetometers usually operate alternatively, to avoid the magnetic rods\u27 noise effect on magnetometers measurements. Because of that, there will be no control authority over the spacecraft during the magnetometer measurement period. Hence longer maneuver times are usually experienced. In this study, a control scheme that enables the extension of the magnetic rods’ activation time is developed, regardless of the attitude control law. The key concept is replacing the real magnetic field measurement by a pseudo measurement, which is computed based on other sensors measurements. By applying a known command to the spacecraft and measuring the spacecraft response, it is possible to compute the ambient magnetic field around the spacecraft. The system mathematical singularity is solved using the Tikhonov regularization approach. Another developed approach estimates the magnetic field, using a relatively simple and fast dynamic model inside a Multiplicative Extended Kalman Filter. A less maneuver time with less power consumption are fulfilled. These control approaches are further validated using real telemetry data from CASSIOPE mission. This dissertation develops a stability analysis for the spacecraft magnetic attitude control, taking into consideration the alternate operation between the magnetic rods and the magnetometers. It is shown that the system stability degrades because of this alternate operation, supporting the proposed approach of extending the operation time of the magnetic rods

The ITASAT CubeSat Development and Design

Because they are inexpensive platforms for satellites, CubeSats have become a low-cost way for universities and even developing countries to have access to space technology. This paper presents the ITASAT design, particularly the Attitude Determination and Control Subsystem, the Onboard Software, and the Assembly, Integration and Testing program. The ITASAT is a 6U CubeSat nano-satellite in development at the Instituto Tecnologico de Aeronautica, in Sao Jose dos Campos, Brazil. The platform and its subsystems will be provided by industry while the payloads are being designed and developed by the principal investigators. The ITASAT Attitude Determination and Control Subsystem will rely on a 3-axis magnetometer, 6 analog cosine sun sensors, 3-axis MEMS gyroscopes, 3 magnetic torque coils, and 3 reaction wheels. The Attitude Determination and Control Subsystem operating modes, control laws, and embedded software are under the responsibility of the Instituto Tecnologico de Aeronautica. A Kalman filter shall be employed to estimate the quaternion attitude and gyroscope biases from sensor measurements. The Attitude Determination and Control Subsystem operating modes are the nominal mode, with geocentric pointing attitude control and the stabilization mode, in which only the satellite angular velocity is controlled. The nominal mode will be split into 2 sub-modes: reaction wheel control plus magnetic wheel desaturation and 3-axis magnetic attitude control. Simulation results have shown that the attitude can be controlled with 1-degree accuracy in nominal mode with the reaction wheels, but these errors grow as much as 20 degrees or higher with the 3-axis magnetic control.Inst Nacl Pesquisas Espaciais Engn, Sao Jose Dos Campos, SP, BrazilTecnol Espaciais Div Mecan Espacial Control, Sao Jose Dos Campos, SP, BrazilInst Tecnol Aeronaut, Dept Ciencia & Tecnol Aeroespacial, Div Engn Aeronaut, Praca Marechal Eduardo Gomes 50 Vila Acacias, BR-12230901 Sao Jose Dos Campos, SP, BrazilUniv Fed Sao Paulo, Inst Ciencia & Tecnol, Dept Ciencia Computacao, Sao Jose Dos Campos, SP, BrazilUniv Fed Sao Paulo, Inst Ciencia & Tecnol, Dept Engn Computacao, Sao Jose Dos Campos, SP, BrazilInst Tecnol Aeronaut, Dept Ciencia & Tecnol Aeroespacial, Div Engn Eletr & Computacao, Sao Jose Dos Campos, SP, BrazilUniv Fed Sao Paulo, Inst Ciencia & Tecnol, Dept Engn Computacao, Sao Jose Dos Campos, SP, BrazilWeb of Scienc

1999 Flight Mechanics Symposium

This conference publication includes papers and abstracts presented at the Flight Mechanics Symposium held on May 18-20, 1999. Sponsored by the Guidance, Navigation and Control Center of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers

Hubble Space Telescope Reduced-Gyro Control Law Design, Implementation, and On-Orbit Performance

Following gyro failures in April 2001 and April 2003, HST Pointing Control System engineers designed reduced-gyro control laws to extend the spacecraft science mission. The Two-Gyro Science (TGS) and One-Gyro Science (OGS) control laws were designed and implemented using magnetometers, star trackers, and Fine Guidance Sensors in succession to control vehicle rate about the missing gyro axes. Both TGS and OGS have demonstrated on-orbit pointing stability of 7 milli-arcseconds or less, which depends upon the guide star magnitude used by the Fine Guidance Sensor. This paper describes the design, implementation, and on-orbit performance of the TGS and OGS control law fine-pointing modes using Fixed Head Star Trackers and Fine Guidance Sensors, after successfully achieving coarse-pointing control using magnetometers