The CanX-4 and CanX-5 mission currently under development at The University of Toronto Institute for Aerospace Studies Space Flight Laboratory UTIAS/SFL is a challenging formation flying technology demonstration. Its requirements of sub-metre control accuracy have yet to be realized with nanosatellites. Many large technical challenges must be addressed in order to ensure the success of the CanX-4/5 mission. This includes the development of software for an intersatellite communication system, integration and optimization of key formation flying algorithms onto the Payload On-Board Computer as well as the development of a Hardware-In-The-Loop simulator for full on-orbit mission simulations. This thesis will provide background knowledge of the Space Flight Laboratory and its activities, the CanX-4/5 mission, and nally highlight the authors contributions to overcoming each of these technical challenges and ensuring the success of the CanX-4 and CanX-5 mission.

A growing interest in distributed systems of small satellites has recently emerged due to their ability to perform a variety of new mission types, increasing technical capability, and reduced time and cost for development. However, the lack of available and dedicated small launch services currently restricts the establishment of these systems in orbit. Secondary payload launch opportunities and alternative deployment strategies can address the issue of access-to-orbit and support the delivery of the constellation to the correct orbit configuration following launch. Of these deployment strategies, the method of indirect plane separation, which utilises the natural precession of Earth orbits, is particularly applicable to the deployment of small satellite constellations due to the potential to significantly reduce propulsive requirements, albeit at the cost of increased deployment time. A review of satellite constellation design revealed that existing methods and tools are not suitable for the analysis of small satellite constellations and are not equipped to investigate alternative deployment strategies, despite the potential benefits of improved access-to-orbit, reduced system complexity, and reduced cost. To address the identified gaps in the design process, a methodology in which the analysis of small satellite constellation deployment is integrated into the system design framework is presented in this thesis. The corresponding system design-space is subsequently explored using a numerical optimisation method, which aids the identification of effective system designs and promotes the understanding of relationships between the design variables and output objectives. The primary objectives of this methodology are to ensure that the different opportunities for deployment of small satellite constellations are thoroughly examined during the design process and to support the development of improved mission and system designs. The presented methodology is demonstrated using a reduced order framework comprised of an analysis for the deployment of small satellite constellations, preliminary vehicle and propulsion system sizing processes, and system cost estimating relationships. Using this simplified mission design framework, the design space-exploration of three small satellite constellation mission case-studies is performed by application of a multiobjective genetic algorithm. Objectives of time-to-deploy, system mass, and system cost are used to direct the optimisation process and search for the most effective solutions in the system design-space. In order to perform the analysis of constellation deployment by the process of indirect plane separation, a simulation method using a semi-analytical propagation technique and time-varying atmospheric density model was developed and verified by comparison to the actual deployment of the FORMOSAT-3/COSMIC mission. The results of the case-studies presented illustrate the ability of the developed methodology to support the design process for satellite constellations and enable the identification of promising and improved system architectures for further development. Moreover, through the enumeration and quantification of the system design-space and tradespace, the methodology is shown to support the identification of relationships and trends between the design variables and selected output objectives, increasing the knowledge available to the system design team during the design process.

Thesis is submitted in fulfilment of the requirements for the degree Master of Technology: Electrical Engineering in the Faculty of Engineering at the Cape Peninsula University of Technology 2013 / The CubeSat standard has various engineering challenges due to its small size and surface area. The challenge is to incorporate a large amount of technology into a form factor no bigger than 10cm3. This research project investigates the space environment, solar cells, secondary sources of power, and Field-Programmable-Gate-Array (FPGA) technology in order to address the size, weight and power challenges presented by the CubeSat standard. As FPGAs have not yet been utilised in this particular sub-system as the main controller, this research investigates whether or not the implementation of an FPGA-based electronic power supply sub-system will optimise its functionality by overcoming these size weight and power challenges. The SmartFusion FPGA was chosen due to its analogue front end which can reduce the number of peripheral components required by such complex systems. Various maximum power point tracking algorithms were studied and it was determined that the perturb-and-observe maximum power point tracking algorithm best suits the design constraints, as it only requires the measurement of either solar cell voltage or solar cell current, thus further decreasing the component count. The SmartFusion FPGA analogue compute engine allows for increased performance of the perturb-and-observe algorithm implemented on the microcontroller sub-system as it allows for the offloading of many repetitive calculations. A VHDL implementation of the pulse-width-modulator was developed in order to produce the various changes in duty cycle produced by the perturb-and-observe algorithm. The aim of this research project was achieved through the development and testing of a nano-satellite power system prototype using the SmartFusion FPGA from Microsemi with a decreased number of peripheral circuits. Maximum power point was achieved in 347ms at worst case with a 55% decrease in power consumption from the estimated 330mW as indicated in the power budget. The SmartFusion FPGA consumes only a worst case of 148.93mW. It was found that the unique features of the SmartFusion FPGA do in fact address the size weight and power constraints of the CubeSat standard within this sub-system.

Among the many possible applications of laser ablation one of the more recent taken is related to nanosatellites propulsion. The study of Laser Ablation Propulsion (LAP) requires research activity on different fields like high power pulsed lasers, laser ablation itself, because it is still a problem to relate the well known mechanisms to impulse generation, and finally materials that represent the fuel in LAP. This thesis presents a research activity on LAP from its very beginning, with the development of an experimental apparatus to measure laser generated impulse and the first results on metals and polymers that paves the way to the development of future LAP materials. Chapter 1 presents an overview of the actual situation of space economy and its recent fast evolution that led in the last years to the exploitation of space for many different applications, also by private companies. The so called New Space Economy is the background on which LAP develops, as an attractive propulsion technique for nano satellites, nowadays extremely diffused in all kind of space missions, and as a possible solution for the space debris problem. In this Chapter typical results obtained in LAP are also reviewed and compared with other solutions both for space debris and propulsion, in order to obtain a better image of its applicability range. Chapter 2 deals with laser ablation. Initially the parameters that play a role in laser ablation are discussed, in particular those related to the laser source like wavelength, pulse duration and repetition rate, to give an overview of the experimental conditions involved. Then general phenomenological observations on laser ablation are presented and related to the physical mechanisms involved, both in the case of metals and polymers, highlighting the main differences between these two classes of materials. The experimental part of this thesis starts in Chapter 3, with the description of the experimental apparatus developed to measure the laser generated mechanical impulses in the order of uN s. The different strategies to perform this kind of measurements are reviewed and compared to the one adopted in this work, based on a ballistic pendulum, and main advantages and problems are discussed. A technical description of the apparatus is given, focusing in particular on all the precautions that have been taken in order to let the pendulum operate in as ideal as possible conditions. The measurement procedure developed during this work is then described in detail, by discussing data analysis and showing some examples. Chapter 4 also deals with the development of the apparatus, in particular for what concerns the estimation of the laser energy density that reaches the target material (fluence), a fundamental parameter for LAP measurements. Some measurements on metals are also presented here in order to discuss some features related to the measurements of some common LAP parameters. Chapter 5 and 6 deal with LAP using polymers, and in particular with experiments devoted to the understanding of material properties that mainly affect LAP performances. The starting material chosen for these experiments is poly(vinyl chloride)(PVC), a benchmark in LAP experiments. Chapter 5 compares localized or uniform laser absorption by PVC, that can be obtained respectively by including carbon nanoparticles in the polymer matrix or by mixing PVC with an absorbing polymer (poly(styrene sulfonate)). The comparison is carried out from the optical an thermodynamical point of view, along with impulse generation. Specific ablation mechanisms are also discussed, showing that a localized absorption of laser radiation is more energetically efficient for impulse generation. Chapter 6 then continues the work on PVC containing nanoparticles, investigating the role of their size, morphology and concentration in laser ablation and in impulse generation. Both commercial and green produced carbon nanoparticles are used for these experiments showing that, at least in the considered size range, the only parameter that affects laser ablation is the number density of absorption centers in the polymer matrix, and not size or morphology. This points the direction to follow in the development of a polymeric material for LAP applications. Some open problems and future works are presented in Chapter 7. Effects on impulse generated by irradiating multiple times the same region are discussed, showing opposite behaviours between metals and polymers, for which still there is not a clear explanation. Then experimental issues and some results on specific impulse measurements are presented, and difficulties related to this measurement in metals briefly discussed. Finally laser ablation in a confined geometry is considered as an attracting technique to enhance impulse generation. And some results on PVC are shown. As a conclusion, main results obtained in this thesis are highlighted, and possible future research activities, developments and perspectives are discussed.

The low size and budget of typical nanosatellite missions limit the available sensors for attitude estimation. Relatively high noise MEMS gyroscopes often must be employed when accurate knowledge of the spacecraft’s angular velocity is necessary for attitude determination and control. This thesis derived and tested in simulation the “Virtual Gyroscope” algorithm, which replaced a standard gyroscope with an array of spatially distributed accelerometers for a 1U CubeSat mission. A MEMS accelerometer model was developed and validated using Root Allan Variance, and the Virtual Gyroscope was tested both in the open loop configuration and as a replacement for a gyroscope in a Multiplicative Extended Kalman Filter. It was found that the quality of the Virtual Gyroscope’s rate measurement improved with a larger and higher quality array, but the error in the estimate was very large. The low signal-to-noise ratio and the unknown bias in the accelerometers caused the angular velocity estimate from the accelerometer array to be too poor for use in the propagation step of the Kalman filter. The Kalman filter performed better with attitude measurements alone than with the Virtual Gyroscope, even when the attitude were delivered at a low rate with added noise. Overall, the current Virtual Gyroscope algorithm that is presented in this thesis is not suitable to replace a MEMS gyroscope in a nanosatellite mission, although there is room for future improvements using bias prediction for the individual accelerometers in the array.

Thesis (MScEng (Mechanical and Mechatronic Engineering))--University of Stellenbosch, 2009. / This thesis describes the development of a vacuum arc thruster (VAT) to be used as a potential low mass (< 500 g), low power (< 5–10W) propulsion system for nanosatellites. The thruster uses a high voltage capacitive circuit to initiate and power the arc process with a 400 ns high current (150–800A) pulse. A one-dimensional steady state analyticalmodel describing the cathode region of the vacuum arc was developed. The model made use of mass and energy balances at the sheath region and cathode surface respectively to predict key quantities such as thrust, ion velocity, ion-to-arc current ratio and erosion rate. Predicted results were shown to be within the limits of reported literature (∼63 μN/A, 26.12 km/s, 0.077 and 110 μg/C respectively). A sensitivity analysis of the analytical model found that a high electric field in the cathode region impedes and decelerates ion flow, which is used for thrust. This was confirmed experimentally for thrust values at arc voltages greater than 2000 V. Both direct and indirect means of measuring thrust were achieved by using a deflecting cantilever beam and an ion collector system, respectively. The transient response of the cantilever beam to impulsive thrust was analytically modeled, whilst the ion current was found by measuring the current induced on a plate subject to ion bombardment. Knowledge of the ion current density distribution was successfully used to approximate the effective normal thrust vector. Direct and indirect thrust levels were roughly 140 and 82 μN/A of average arc current, respectively. Measured thrust was found to be higher than predicted thrust due to thrust contributions fromthe ablation of Teflon insulation. The discrepancy is also due to the uncertainty in quantifying free parameters in the analytical model such as the fraction of generated ions flowing away from the cathode region. The thrust-topower ratio, specific impulse and efficiency of the vacuum arc thruster at an average arc current of 200 A was measured to be 0.6 μN/W, 160 s and 0.05 %, respectively. A thruster performance analysis and specification showed that the VAT is capable of achieving specific orbital and slew manoeuvres within a constant 5–10 W average power. It was concluded that thruster performance could be improved by using a two-stage arc circuit consisting of a high voltage, low current, short pulse trigger and a low voltage, high current, long pulse driver.

Dans cette thèse, le problème du contrôle d’attitude de satellites agiles à l’aide de grappes d’actionneurs gyroscopiques (AGs) est considéré et plus particulièrement son application au contrôle de micro/nanosatellites (10-100 kg). Des outils d’analyse topologique sont tout d'abord développés. La comparaison de différentes configurations de grappes justifie le choix d'une géométrie pyramidale à six actionneurs gyroscopiques. Des analyses plus approfondies de cette grappe, avec et sans cas de panne, permettent de déduire des contraintes que la loi de pilotage doit vérifier pour être adaptée à ce système. Pour y répondre, après analyse de la littérature, une nouvelle structure de loi de pilotage ainsi qu’une formulation différente des équations cinématiques sont développées. Cette structure est basée sur l’algorithme du filtre de Kalman étendu. Elle a pour avantages de répondre aux exigences en termes de calcul temps réel au bord des satellites, de flexibilité sur la gestion des contraintes et de facilité d’adaptation en cas de pannes. En outre, une procédure de génération de boucle de commande, englobant la loi de pilotage et un contrôleur robuste du système, est proposée. La généralisation de cette boucle de commande est illustrée sur des bras manipulateurs à base fixe et spatiaux.En parallèle, l’étude du passage des singularités internes intraversables mène à une nouvelle stratégie d’évitement de ces singularités. Des simulations sur des modèles de satellites représentatifs illustrent les résultats. La grappe d’actionneurs et la boucle de commande développées seront testées dans le cadre d’une expérimentation en microgravité. / In this thesis, the attitude control problem for agile satellites with control moment gyro (CMG) clusters is studied. In particular, the problem applies to micro/nanosatellites (10-100kg). First, numerical tools are developed to analyse the compatibility of various cluster configurations with the nanosatellite constraints. The pyramidal six-CMG cluster is then selected. This cluster topology is analysed in depth, with and without actuator failures. Constraints on the development of a steering law adapted to our system are deduced. Among them, the singularity avoidance issue is emphasised. To meet the requirements, an analysis of the literature is carried out. Then, a new steering law structure and a different formulation of the kinematic equations are developed. This structure is based on the extended Kalman filter algorithm. It meets the requirements because it can be calculated in real-time onboard satellites, constraints imposed on the system are handled with flexibility and it is easily adaptable in case of actuator failures. In addition, a procedure to generate the control loop is proposed, containing a robust controller. The generalisation of this control loop is shown on space and fixed-base manipulator arms. Furthermore, the study of the internal elliptic singularities in CMG clusters leads to a new singularity avoidance strategy. Software simulations on highly representative simulators show the results of the steering law in various actuator failure cases. The CMG cluster and the control loop will be tested in a parabolic flight campaign, and the development of this experiment is detailed in this thesis.

The analysis, verification and emulation of attitude control hardware for nanosatellite spacecraft is described. The overall focus is on hardware that pertains to a multitude of missions currently under development at the University of Toronto Institute for Aerospace Studies - Space Flight Laboratory. The requirements for these missions push the boundaries of what is currently the accepted performance level of attitude control hardware. These new performance envelopes demand new acceptance test methods which must verify the performance of the attitude control hardware. In particular, reaction wheel and hysteresis rod actuators are the focus. Results of acceptance testing are further employed in post spacecraft integration for hardware emulation. This provides for a reduced mission cost as a function of reduced spare hardware. The overall approach provides a method of acceptance testing to new performance envelopes with the benefit of cost reduction with hardware emulation for simulations during post integration.

The analysis, verification and emulation of attitude control hardware for nanosatellite spacecraft is described. The overall focus is on hardware that pertains to a multitude of missions currently under development at the University of Toronto Institute for Aerospace Studies - Space Flight Laboratory. The requirements for these missions push the boundaries of what is currently the accepted performance level of attitude control hardware. These new performance envelopes demand new acceptance test methods which must verify the performance of the attitude control hardware. In particular, reaction wheel and hysteresis rod actuators are the focus. Results of acceptance testing are further employed in post spacecraft integration for hardware emulation. This provides for a reduced mission cost as a function of reduced spare hardware. The overall approach provides a method of acceptance testing to new performance envelopes with the benefit of cost reduction with hardware emulation for simulations during post integration.

Thesis (MEng)-- Stellenbosch University, 2013. / ENGLISH ABSTRACT: The next generation of CubeSats will require accurate attitude knowledge throughout orbit for advanced science payloads and high gain antennas. A star tracker can provide the required performance, but star trackers have traditionally been too large, expensive and power hungry to be included on a CubeSat. The aim of this project is to develop and demonstrate a CubeSat compatible star tracker. Subsystems from two other CubeSat components, CubeSense and CubeComputer, were combined with a sensitive, commercial image sensor and low-light lens to produce one of the smallest star trackers in existence. Algorithms for star detection, matching and attitude determination were investigated and implemented on the embedded system. The resultant star tracker, named CubeStar, can operate fully autonomously, outputting attitude estimates at a rate of 1 Hz. An engineering model was completed and demonstrated an accuracy of better than 0.01 degrees during night sky tests. / AFRIKAANSE OPSOMMING: Die volgende generasie van CubeSats sal akkurate orientasie kennis vereis gedurende 'n volle omwentelling van die aarde. 'n Sterkamera kan die vereiste prestasie verskaf, maar sterkameras is tradisioneel te groot, duur en krag intensief om ingesluit te word aanboord 'n CubeSat. Die doel van hierdie projek is om 'n CubeSat sterkamera te ontwikkel en te demonstreer. Substelsels van twee ander CubeSat komponente, CubeSense en CubeComputer, was gekombineer met 'n sensitiewe kommersiële beeldsensor en 'n lae-lig lens om een van die kleinste sterkameras op die mark te produseer. Algoritmes vir die ster opsporing, identi kasie en orientasie bepaling is ondersoek en geïmplementeer op die ingebedde stelsel. Die gevolglike sterkamera, genaamd CubeStar, kan ten volle outonoom orientasie afskattings lewer teen 'n tempo van 1 Hz. 'n Ingenieursmodel is voltooi en 'n akkuraatheid van beter as 0.01 grade is gedemonstreer.