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

Precise Relative Orbit Determination of Low Earth Orbit Formation Flights using GPS Pseudorange and Carrier-Phase Measurements

Formation flying is emerging as an important technology on achieving the tight mission requirements of imaging and remote sensing systems, especially radio interferometry and synthetic aperture radar (SAR) applications. A higher absolute and relative position and orbit knowledge is always sought in these kinds of applications. Such requirements can be met to a large extent by manipulation of GPS data. Carrier-phase Differential GPS (CDGPS) measurements can also be used to further increase the accuracy in relative position and orbit determination dramatically. Using a geometric model has a clear advantage of generality and wide applicability, independent of complex dynamic models for different types of platforms. Hence, the proposed approach uses input from GPS receiver on the master satellite and pseudorange based absolute position estimates from the slave satellites. In addition, single-difference (SD) phase measurements between the master and the slave satellites are also required, which provide very accurate relative distance information. SD information is input into a Kalman filter to determine the relative orbits within the formation to a higher precision. In this paper, we present a geometrical approach to relative orbit determination and present an algorithm for the refinement of position estimates through combining carrier-phase and pseudorange data

A Study into the Method of Precise Orbit Determination of a HEO Orbiter by GPS and Accelerometer

In the present day, orbit determination by Global Positioning System (GPS) is not unusual. Especially for low-cost small satellites, position determination by an on-board GPS receiver provides a cheap, reliable and precise method. However, the original purpose of GPS is for ground users, so the transmissions from all of the GPS satellites are directed toward the Earth s surface. Hence there are some restrictions for users above the GPS constellation to detect those signals. On the other hand, a desire for precise orbit determination for users in orbits higher than GPS constellation exists. For example, the next Japanese Very Long Baseline Interferometry (VLBI) mission "ASTRO-G" is trying to determine its orbit in an accuracy of a few centimeters at apogee. The use of GPS is essential for such ultra accurate orbit determination. This study aims to construct a method for precise orbit determination for such high orbit users, especially in High Elliptical Orbits (HEOs). There are several approaches for this objective. In this study, a hybrid method with GPS and an accelerometer is chosen. Basically, while the position cannot be determined by an on-board GPS receiver or other Range and Range Rate (RARR) method, all we can do to estimate the user satellite s position is to propagate the orbit along with the force model, which is not perfectly correct. However if it has an accelerometer (ACC), the coefficients of the air drag and the solar radiation pressure applied to the user satellite can be updated and then the propagation along with the "updated" force model can improve the fitting accuracy of the user satellite s orbit. In this study, it is assumed to use an accelerometer available in the present market. The effects by a bias error of an accelerometer will also be discussed in this paper

Autonomous Control System for Precise Orbit Maintenance

In this paper, we describe a closed-loop autonomous control system that enables orbit operations to be performed without the need of any ground segment. The growing availability of GPS receivers on satellites provides an excellent means for autonomous orbit determination and our work builds upon previous work on orbit determination algorithms developed here at Surrey. The orbit is described using a set of epicycle parameters which provide an analytic model of LEO orbits. The parameters in this model are estimated onboard the satellite using a Kalman filter. We describe an enhancement to this software which provides both control as well as estimation of the orbit parameters and a discussion of how atmospheric drag has been included in the model. The goal of the control part of the software is to ensure that the orbital altitude of the satellite never falls outside of a prescribed window due to drag. We present results of the orbit maintenance software which has been successfully running on Surrey\u27s minisatellite UoSat-12. This satellite is in a 650 km altitude orbit at inclination 64.57o . The satellite has been manoeuvred into a repeat ground track orbit so that the satellite repeats its ground track every 7 days. The orbit maintenance software then attempts to maintain the satellite in its resonant orbit, and also to slowly manoeuvre the satellite into a frozen orbit so that the altitude at each pass does not vary

Optimal Combined ReactionWheel Momentum Management for Earth-Pointing Satellites

Abstract: The optimal controllers for the management of 3-axis reaction-wheel momentum of rigid Earth-pointing satellites are analyzed in detail using magnetorquers and/or thrusters. Especially, two novel, optimal combined control schemes are proposed in order to achieve rapid, propellant-saving reaction wheel momentum dumping control by employing magnetorquers and thrusters. Finally, simulation results are presented to demonstrate the superiority of these algorithms. These two combined algorithms could easily be applied in real-time onboard an LEO Earth-pointing satellite. I

Optimal Q-laws via reinforcement learning with guaranteed stability

Closed-loop feedback-driven control laws can be used to solve low-thrust many-revolution trajectory design and guidance problems with minimal computational cost. Lyapunov-based control laws offer the benefits of increased stability whilst their optimality can be increased by tuning their parameters. In this paper, a reinforcement learning framework is used to make the parameters of the Lyapunov-based Q-law state-dependent, increasing its optimality. The Jacobian of these state-dependent parameters is available analytically and, unlike in other optimisation approaches, can be used to enforce stability throughout the transfer. The results focus on GTO–GEO and LEO–GEO transfers in Keplerian dynamics, including the effects of eclipses. The impact of the network architecture on the behaviour is investigated for both time- and mass-optimal transfers. Robustness to navigation errors and thruster misalignment is demonstrated using Monte Carlo analyses. The resulting approach offers potential for on-board autonomous transfers and orbit reconfiguration

Commissioning of a Small Satellite Constellation - Methods and Lessons Learned

The Disaster Monitoring Constellation consists of a series of four spacecraft built by Surrey Satellite Technology Ltd. The aim of the constellation is to be able to perform daily revisits over any point on the Earth such that 32m resolution images can be taken. These images can be used for daily monitoring of natural and man-made disasters. The constellation is made up of four spacecraft. Alsat-1, is an 88 kg microsatellite, which was launched on 28th Nov 2002, for the Algerian national space agency CNTS. It was joined in orbit by three additional microsatellites on 27th September 2003, Nisat-1 (Nigerian national space agency), Bilsat-1 (Tubitak Bilten, Turkey) and UK-DMC (British National Space Centre). In order to exploit small satellites in constellations, a number of key technologies must be utilised. These include precise navigation, propulsion, and coordinated maneuvers planning. This paper describes the spacecraft, the key technologies and how they were used to form the four spacecraft into a working constellation