LEO GPS attitude determination algorithm for a micro-satellite using boom-arm deployed antennas

This paper describes a low earth orbiter micro-satellite attitude determination algorithm using GPS phase and pseudorange data as the only observables. It is designed to run in real-time, at a rate of 10 Hz, on-board the spacecraft, using minimal chip and memory resources. The spacecraft design includes four GPS antennas deployed on boom arms to improve the antenna separations. The boom arms feature smart sensors, from which time-varying deformation data are used to calculate changes in the body-fixed system (BFS) co-ordinates of the attitude antennas. These data are used as input to the attitude algorithm to improve the accuracy of the output. The conventional double-difference phase observation equations have been re-arranged so that the only unknown parameters in the functions (once the ambiguities have been determined) are the spacecraft Euler angles. This greatly increases the redundancy in the mathematical model, and is exploited to enhance the algorithm's ability to trap observations contaminated by unmodelled multipath. This approach has been shown to be successful in identifying phase outliers at the 5–10 mm level. Speed of execution of the program is improved by utilising numerical differentiation of the model equations in the linearisation process. Furthermore, as the number of solve-for parameters is reduced to three by the chosen mathematical model, matrix inversion requirements are minimised. A novel approach to ambiguity resolution and determination of initial estimates of the attitude parameters has been developed utilising a heuristic technique and the known, and time varying, BFS co-ordinates of the antenna array. Algorithm testing is based on a simulation of the micro-satellite trajectory combined with variations in attitude derived from spin-stabilisation and periodic roll and pitch parameters. The trajectory of the spacecraft centre of mass was calculated by numerical integration of a force model using Earth gravity field parameters, third body effects due to the Sun and the Moon, dynamic Earth tide effects (solar and lunar), and a solar radiation pressure model. Frame transformations between J2000 and ITRF97 used the IERS conventions. A similar approach was used to calculate the trajectories of all available GPS satellites during the same period, using initial conditions of position and velocity from IGS precise orbits. RMS differences between the published precise orbit and the integrated satellite positions were at the 5-mm level. Phase observables are derived from these trajectories, biased by simulation of receiver and satellite clock errors, cycle slips, random or systematic noise and initial integer ambiguities. In the actual simulation of the attitude determination process in orbit, GPS satellite positions are calculated using broadcast ephemerides. The results show that the aim of 0.05° (two sigma) attitude precision can be met provided that the phase noise can be reduced to the level of 1–2 mm. Attitude precision was found to vary strongly with constellation geometry, which can change quite rapidly depending on the variations in spacecraft attitude. The redundancy in the mathematical model was found to be very effective in trapping and isolating cycle slips to the double difference observations that are contaminated. This allows for the possibility of correcting for cycle slips without full recourse to the ambiguity resolution algorithm. Electronic Publication