An autonomous, real-time orbit determination system was developed within thiswork for the next iteration of the University of W¨urzburg’s CubeSat programme.The algorithm only made use of magnetometer and sun sensors, which already wereimplemented on UWE-3, the third satellite in the programme. Previous developedsystems used the same approach, however the unique aspect in this work is thatthe algorithm was implemented using equinoctial elements.A Runge-Kutta-4 integrator propagated the orbit position using the orbit dynamicsunder the consideration of J2-perturbations. Afterwards, an Extended KalmanFilter corrected the position through processing the two measurements.The algorithm was then tested under multiple conditions. At first, a two weekstability test was conducted using simulated data, followed by a test with recordedsatellite data. These have shown a mean error of 13.2 km and 12.6 km respectively.Lastly, the algorithm was translated in to C and evaluated on a micro-controller.

Due to the absence of any satellite-based global positioning system on Mars, the Mars Exploration Rovers commonly track position changes of the vehicle using a technique called visual odometry (VO), where updated rover poses are determined by tracking keypoints between stereo image pairs. Unfortunately, the error of VO grows super-linearly with the distance traveled, primarily due to the contribution of orientation error. This thesis outlines a novel approach incorporating sun sensor and inclinometer measurements directly into the VO pipeline, utilizing absolute orientation information to reduce the error growth of the motion estimate. These additional measurements have very low computation, power, and mass requirements, providing a localization improvement at nearly negligible cost. The mathematical formulation of this approach is described in detail, and extensive results are presented from experimental trials utilizing data collected during a 10 kilometre traversal of a Mars analogue site on Devon Island in the Canadian High Arctic.

Due to the absence of any satellite-based global positioning system on Mars, the Mars Exploration Rovers commonly track position changes of the vehicle using a technique called visual odometry (VO), where updated rover poses are determined by tracking keypoints between stereo image pairs. Unfortunately, the error of VO grows super-linearly with the distance traveled, primarily due to the contribution of orientation error. This thesis outlines a novel approach incorporating sun sensor and inclinometer measurements directly into the VO pipeline, utilizing absolute orientation information to reduce the error growth of the motion estimate. These additional measurements have very low computation, power, and mass requirements, providing a localization improvement at nearly negligible cost. The mathematical formulation of this approach is described in detail, and extensive results are presented from experimental trials utilizing data collected during a 10 kilometre traversal of a Mars analogue site on Devon Island in the Canadian High Arctic.

The Generic Nanosatellite Bus (GNB) is a spacecraft platform designed to accommodate the integration of diverse payloads in a common housing of supporting components. The development of the GNB at the Space Flight Laboratory (SFL) under the Canadian Advanced Nanospace eXperiment (CanX) program provides accelerated access to space while reducing non-recurring engineering (NRE) costs. The work presented herein details the development of the attitude determination and control subsystem (ADCS) of the GNB. Specific work on magnetorquer coil assembly, integration, and testing (AIT) and reaction wheel testing is included. The embedded software development and unit-level testing of the GNB sun sensors are discussed. The characterization of the AeroAstro star tracker is also a major focus, with procedures and results presented here. Hardware models were developed and incorporated into SFL's in-house high-fidelity attitude dynamics and control simulation environment. This work focuses on specific contributions to the CanX-3, CanX-4&5, and AISSat-1 nanosatellite missions.

The Generic Nanosatellite Bus (GNB) is a spacecraft platform designed to accommodate the integration of diverse payloads in a common housing of supporting components. The development of the GNB at the Space Flight Laboratory (SFL) under the Canadian Advanced Nanospace eXperiment (CanX) program provides accelerated access to space while reducing non-recurring engineering (NRE) costs. The work presented herein details the development of the attitude determination and control subsystem (ADCS) of the GNB. Specific work on magnetorquer coil assembly, integration, and testing (AIT) and reaction wheel testing is included. The embedded software development and unit-level testing of the GNB sun sensors are discussed. The characterization of the AeroAstro star tracker is also a major focus, with procedures and results presented here. Hardware models were developed and incorporated into SFL's in-house high-fidelity attitude dynamics and control simulation environment. This work focuses on specific contributions to the CanX-3, CanX-4&5, and AISSat-1 nanosatellite missions.

Observing the relative state of two space vehicles has been an active field of research since the earliest attempts at space rendezvous and docking during the 1960's. Several techniques have successfully been employed by several space agencies and the importance of these systems has been repeatedly demonstrated during the on-orbit assembly and continuous re-supply of the International Space Station. More recent efforts are focused on technologies that can enable fully automated navigation and control of space vehicles. Technologies which have previously been investigated or are actively researched include Video Guidance Systems (VGS), Light Detection and Ranging (LIDAR), RADAR, Differential GPS (DGPS) and Visual Navigation Systems. The proposed system leverages the theoretical foundation which has been advanced in the development of VisNav, invented at Texas A & M University, and the miniaturized commercially available Northstar sensor from Evolution Robotics. The dissertation first surveys contemporary technology, followed by an analytical investigation of the coarse sun sensor and errors associated with utilizing it in the near-field. Next, the commercial Northstar sensor is investigated, utilizing fundamentals to generate a theoretical model of its behavior, followed by the development of an experiment for the purpose of investigating and characterizing the sensor's performance. Experimental results are then presented and compared with a numerical simulation of a single-sensor system performance. A case study evaluating a two sensor implementation is presented evaluating the proposed system's performance in a multisensor configuration. The initial theoretical analysis relied on use of the cosine model, which proved inadequate in fully capturing the response of the coarse sun sensor. Fresenel effects were identified as a significant source of unmodeled sensor behavior and subsequently incorporated into the model. Additionally, near-field effects were studied and modeled. The near-field effects of significance include: unequal incidence angle, unequal incidence power, and non-uniform radiated power. It was found that the sensor displayed inherent instabilities in the 0.3 degree range. However, it was also shown that the sensor could be calibrated to this level. Methods for accomplishing calibration of the sensor in the near-field were introduced and feasibility of achieving better than 1 cm and 1 degree relative position and attitude accuracy in close proximity, even on a small satellite platform, was determined.

Experimental testing of attitude determination systems still plays an important role, despite increasing use of simulations. Testing provides a means to numerically quantify system performance, give confidence in the models and methods, and also discover and compensate for unexpected behaviours and interactions with the attitude determination system. The usefulness of the test results is dependent on an understanding of the uncertainties that contribute to the attitude error. With this understanding, the significance of the results can be assessed, and efforts to reduce attitude errors can be directed appropriately. The work of this thesis is to gain a quantitative understanding of the uncertainties that impact the attitude error of low cost spinning spacecraft using COTS camera (as Sun sensor) and MEMS magnetometer. The sensors were calibrated and the uncertainties in these calibrations were quantified, then propagated through the Triad method to uncertainties in the attitude. It was found that most systematic errors were reduced to negligible levels, except those due to timing latencies. Attitude errors achieved in the laboratory with the experimental setup were around 0.14 degrees (3σ) using either the Triad, q-method or Extended Kalman Filter with a gyro for dynamic model replacement. The errors in laboratory were dominated by magnetometer noise. Furthermore, correlated systematic errors had the effect of reducing the attitude error calculated in the laboratory. For an equivalent Sun-mag geometry in orbit, simulation showed that total attitude error would be of the order of 0.77 degrees (3σ). An uncertainty contribution analysis revealed this error was dominated by uncertainties in the inertial magnetic field model. Uncertainties in knowledge of the inertial Sun model, sensor calibration, sensor alignment and sensor noise were shown to be insignificant in comparison.