Trotz der beeindruckenden Fortschritte, die die Röntgen- und Gammastrahlenobservatorien in den letzten Jahrzehnten erzielt haben, ist der Energiebereich zwischen 200 keV und 50 MeV nach wie vor kaum erforscht. Diese Lücke, die in der Literatur oft als ``MeV-Lücke'' bezeichnet wird, ist nicht auf einen Mangel an überzeugender Wissenschaft zurückzuführen, sondern auf technische Herausforderungen und Nachweisschwierigkeiten, die mit MeV-Beobachtungen einhergehen. COMPTEL an Bord von CGRO (1991-2000) war das letzte Teleskop, das eine vollständige Durchmusterung des MeV-Himmels mit einer relativ bescheidenen Empfindlichkeit durchführte. Für die Zukunft sind zahlreiche Missionen vorgeschlagen worden, insbesondere AMEGO, die die Leistung von COMPTEL um mindestens eine Größenordnung verbessern sollen. Der Zeitrahmen für die Entwicklung, den Aufbau und den Start solch großer Missionen beträgt jedoch etwa 10 Jahre und ist mit erheblichen Kosten verbunden. In diesem Szenario könnte ein viel kleinerer Satellit, der sich der neuen Welle von schnellen, relativ kostengünstigen Weltraumforschungsmissionen anschließt, die durch CubeSats ermöglicht werden, in kürzerer Zeit rentabel sein. In dieser Arbeit werden die Verfügbarkeit und die Leistung eines Compton-Teleskops auf der Grundlage des CubeSat-Standards, genannt MeVCube, untersucht. Die Auswirkungen der Materialwahl und verschiedener CubeSat-Nutzlasten wurden durch Simulationen bewertet. Trotz der begrenzten Größe kann selbst ein kleines Teleskop, das auf einem CubeSat fliegt, den Energiebereich von Hunderten von keV bis zu einigen MeV mit einer Empfindlichkeit abdecken, die mit der der letzten Generation von Großmissionen wie COMPTEL und INTEGRAL vergleichbar ist. Es wurden auch experimentelle Messungen an Cadmium-Zink-Tellurid-Halbleiterdetektoren und einer für den Weltraumbetrieb geeigneten Ausleseelektronik mit geringem Stromverbrauch durchgeführt. / Despite the impressive progresses achieved both by X-ray and gamma-ray observatories in the last decades, the energy range between 200 keV and 50 MeV remains poorly explored. This gap in coverage, often referred in literature as the ``MeV gap'', is not due to lack of compelling science, but instead to technical challenges and detection difficulties that comes with MeV observations. COMPTEL, on-board CGRO (1991-2000), was the last telescope to accomplish a complete survey of the MeV-sky with a relatively modest sensitivity. Many missions have been proposed for the future, most notably AMEGO, aiming to improve COMPTEL's performance by at least one order of magnitude. However, the timescale for development, assembly and launch of such large missions is around 10 years, with substantial costs. Looking at this scenario, a much smaller satellite, joining the new wave of rapid, relatively inexpensive space science missions enabled by CubeSats, may be profitable on a shorter time-scale. This thesis evaluates the availability and performance of a Compton telescope based on the CubeSat standard, named MeVCube. The impact of material choice and different CubeSat payloads has been evaluated through simulations. Despite the limited size, even a small telescope flying on a CubeSat can cover the energy range from hundreds of keV up to few MeVs with a sensitivity comparable to that of the last generation of large-scale missions like COMPTEL and INTEGRAL. Experimental measurements on Cadmium-Zinc-Telluride semiconductor detectors and low-power read-out electronics suitable for space operation have been performed as well.

Small satellites and CubeSats have established themselves within the aerospace community because of their low cost and high return on investment. Many CubeSats are developed in a short time frame and often leverage commercial off the shelf components for quick turnaround missions. With regard to the Electrical Power System, commercially available products typically use a centralized architecture. However, a centralized architecture is not reusable, since missions that require additional solar arrays or batteries would necessitate a redesign of the power system. With the range of CubeSat sizes and mission goals, it is obvious that a one-size-fits-all solution is not appropriate. This thesis details a reusable and scalable power system architecture applicable to a variety of missions. Reusability is achieved by using common building blocks or "modules," where the same modules can be used between missions. Scalability is achieved by not limiting the number of modules that can be connected together—more modules can be added as needed. In this system, solar arrays and battery units connect directly to a common bus, supplying an unregulated voltage to each subsystem. These subsystems then regulate the bus voltage to their individual needs. The power system also features direct energy transfer and solar-only operation.

Future spacecraft missions are trending towards the use of distributed systems or fractionated spacecraft. Initiatives such as DARPA's System F6 are encouraging the satellite community to explore the realm of collaborative spacecraft teams in order to achieve lower cost, lower risk, and greater data value over the conventional monoliths in LEO today. Extensive research has been and is being conducted indicating the advantages of distributed spacecraft systems in terms of both capability and cost. Enabling collaborative behaviors among teams or formations of pico-satellites requires technology development in several subsystem areas including attitude determination and control subsystems, orbit determination and maintenance capabilities, as well as a means to maintain accurate knowledge of team members' position and attitude. All of these technology developments desire improvements (more specifically, decreases) in mass and power requirements in order to fit on pico-satellite platforms such as the CubeSat. In this thesis a solution for the last technology development area aforementioned is presented. Accurate knowledge of each spacecraft's state in a formation, beyond improving collision avoidance, provides a means to best schedule sensor data gathering, thereby increasing power budget efficiency. Our solution is composed of multiple software and hardware components. First, finely-tuned flight system software for the maintaining of state knowledge through equations of motion propagation is developed. Additional software, including an extended Kalman filter implementation, and commercially available hardware components provide a means for on-board determination of both orbit and attitude. Lastly, an inter-satellite communication message structure and protocol enable the updating of position and attitude, as required, among team members. This messaging structure additionally provides a means for payload sensor and telemetry data sharing. In order to satisfy the needs of many different missions, the software has the flexibility to vary the limits of accuracy on the knowledge of team member position, velocity, and attitude. Such flexibility provides power savings for simpler applications while still enabling missions with the need of finer accuracy knowledge of the distributed team's state. Simulation results are presented indicating the accuracy and efficiency of formation structure knowledge through incorporation of the described solution. More importantly, results indicate the collaborative module's ability to maintain formation knowledge within bounds prescribed by a user. Simulation has included hardware-in-the-loop setups utilizing an S-band transceiver. Two "satellites" (computers setup with S-band transceivers and running the software components of the collaborative module) are provided GPS inputs comparable to the outputs provided from commercial hardware; this partial hardware-in-the-loop setup demonstrates the overall capabilities of the collaborative module. Details on each component of the module are provided. Although the module is designed with the 3U CubeSat framework as the initial demonstration platform, it is easily extendable onto other small satellite platforms. By using this collaborative module as a base, future work can build upon it with attitude control, orbit or formation control, and additional capabilities with the end goal of achieving autonomous clusters of small spacecraft.

Thesis (MTech (Electrical Engineering))--Cape Peninsula University of Technology. / Front-end power converters for nanosatellite applications demand better performance in accurate reference tracking because of the wide-range input voltage of the solar panels. The very tight output voltage requirements demand a robust, reliable, and high-efficiency converter. The control of such a converter is very complex and time consuming to design. Two commonly used control modes are current and voltage control. The design and implementation of a voltage controller for DC–DC power converter is simpler but compared to current mode controller, does not do provide for overcurrent protection. A single-ended primary inductance converter (SEPIC) was selected for this research work because of its ability to buck or boost the input voltage coupled with the ability to provide noninverting polarity with respect to the input voltage. Parameter values for the converter studied are used to analyse and design both the voltage and the current mode controllers for the nanosatellite front-end power converter. Output voltage reference tracking with step and ramp changes in the input voltage is evaluated in terms of the time taken to reach steady-state after the induced disturbances and either the overshoot or undershoot of the output voltage reference. The design of analogue pulse width modulation (PWM) study was carried out in order to drive the metal-oxide-semiconductor field-effect transistor (MOSFET) switch. For the two controllers, changes in the reference output voltage in response to load changes are also studied. An examination of the effects of solar radiation on the MOSFET switch was conducted; this switch is the main component of the front-end DC–DC power converter for a nanosatellite. At the more general level the examination also provided information on the response of the semiconductor technology in space application. The overall purpose of studying the MOSFET switch was to investigate the mechanisms that will facilitate its ability of switching ‘on’ and ‘off’ without failure as a result of solar radiation. The effects of solar radiation on MOSFET device in space, has resulted in more malfunctions of these devices in the past five years than over the preceding 40 years.

<div><div><div><p>The advent of nano and micro class satellites has generated new demand for compact and efficient propulsion systems. Traditional propulsion technologies have been miniaturized for the CubeSat platform and new technology solutions have been proposed to address this demand. However, each of these approaches has disadvantages when applied within the context of a CubeSat. One potential low mass and power alternative is Film-Evaporation MEMS Tunable Array (FEMTA) micropropulsion which is capable of generating 150μN of thrust using 0.65W of electrical power and ultra-pure deionized water as propellant. The FEMTA thruster is etched into a 1cm × 1cm × 0.3mm silicon substrate using standard photolithography and microfabrication techniques. Each thruster consists of a 4 μm wide nozzle and platinum resistive heaters. Capillary pressure prevents the water from leaking through the nozzle and the heaters induce film-evaporation at the fluid interface to generate thrust. FEMTA has been in development at Purdue University since 2015 under the NASA SmallSat Technology Partnership Program and is currently on its 5th generation design. While these generations of FEMTA have successfully demonstrated the viability of the propulsion technique under ideal conditions, multiple reliability and performance related issues have been identified. More specifically, high vacuum tests have shown that the current FEMTA design is susceptible to quiescent propellant mass loss due to ice generation and leaking at the nozzle. These mass ejections can limit the lifespan and performance of the thruster and can induce undesired attitude perturbations on the host spacecraft. The purpose of this researchidentify the root causes of the quiescent mass loss mechanims hrough simulation and direct experimentation. Based on the results of these investigations, a next generation design is proposed, fabricated, and tested. Microfabrication was performed at Purdue’s Birck Nanotechnology Center and vacuum and thrust stand tests were performed at the High Vacuum Lab in the Aerospace Sciences Laboratory at Purdue.</p></div></div></div>

The magnetic cleanliness of a spacecraft during magnetic measurements is an important aspect in the design of many space science missions. The adequate reduction or removal of the spacecraft's magnetic disturbance plays a vital role in allowing the ambient magnetic field to be measured with the required accuracy. There are three main approaches to reduce the impact of the spacecraft's magnetic disturbance on the final magnetic measurement, with each approach imposing its own set of changes and constraints on the spacecraft. In turn these changes and constraints introduce additional complexity and cost to the system design. The required changes in the spacecraft's mission profile and configuration also need to be factored in during the design phase of a spacecraft, but cannot be avoided if high quality measurements are desired.  One of these approaches is the use of a magnetic cleanliness programme, and such programmes have a long history of successful use on large satellite missions.CubeSats have become increasingly technically capable and have in recent years begun to undertake scientific missions with challenging sensitivity requirements, including for magnetic measurements. The small size of the CubeSat form factor poses some unique challenges to the implementation of magnetic cleanliness techniques, but are also in increased need of limiting the residual magnetic moment when compared to large satellites. This thesis details the early phases of the magnetic cleanliness programme on the FORESAIL-2 science mission. Nine magnetic cleanliness requirements on the FORESAIL-2 satellite platform were derived from the FORESAIL-2 measurement and instrument requirements. A simple magnetic model was established, the results of which were used to propose a configuration of spacecraft subsystems. The resulting preliminary configuration of subsystems reduced the effective magnetic field of the REPE payload at the sensor by 352pT, 30.1%, when compared to the worst-performing configuration. Subsequently an improved model, utilising RSS analysis, was created. Combined with updated location information for each subsystem, defined using the proposed configuration, this second model yielded an estimated magnetic field of 2710pT at the reference point. The results of the second model were also used to identify the CDE payload and the TT&amp;C subsystem as potentially problematic from a magnetic cleanliness perspective. A list of ferromagnetic materials was compiled, and a total mass of 453.72g of ferromagnetic materials was estimated. The work presented in this thesis is expected to form the basis of FORESAIL-2's continued magnetic cleanliness programme throughout the design and integration phases of the project.  For example, the improved model could be extended to include the estimated magnetic dipole moment's of each subsystem. Additionally, the inventories of materials, currents, and frequencies established as part of this thesis can be maintained throughout the FORESAIL-2's development cycle, and used to update the estimated total magnetic field of the spacecraft platform.

Langmuir probes are extensively employed to study the plasmas in space and laboratory environments. Successful measurements require a comprehensive modeling of both the plasma environment and the probe conditions in the form of current collection models. In this thesis, the surface morphology implications on the probe current collection are investigated. This problem is applied and solved in the context of a CubeSat regime. The first problem that is investigated is the consequence of surface structural variability on the current measurements. A new model for dealing with non-uniformity of the probe surface structure is developed in this paper. This model is applied to analyze the Langmuir probe data from a sounding rocket mission that was subjected to surface structural non-homogeneities. This model would be particularly useful for CubeSat platforms where elaborate probe design procedures are not feasible. The second problem that is investigated is the surface area implications on Langmuir probe measurements. It has been established that surface area ratio of the spacecraft to that of the probe needs to be sufficiently large to make successful plasma measurements. CubeSats would therefore pose a challenge for employing Langmuir-type instruments to study the space plasma. We inspect the feasibility of making plasma measurements using Langmuir probes subjected to CubeSat area constraints. This analysis is done for a forthcoming Utah State University (USU)/Space Dynamics Lab (SDL) CubeSat mission.

As development of CubeSat based architectures increase, methods of deploying constellations of CubeSats are required to increase functionality of future systems. Given their low cost and quickly increasing launch opportunities, large numbers of CubeSats can easily be developed and deployed in orbit. However, as secondary payloads, CubeSats are severely limited in their options for deployment into appropriate constellation geometries. This thesis examines the current methods for deploying cubes and proposes new and efficient geometries using secondary launch opportunities. Due to the current deployment hardware architecture, only the use of different launch opportunities, deployment direction, and deployment timing for individual cubes in a single launch are explored. The deployed constellations are examined for equal separation of Cubes in a single plane and effectiveness of ground coverage of two regions. The regions examined are a large near-equatorial zone and a medium sized high latitude, high population density zone. Results indicate that simple deployment strategies can be utilized to provide significant CubeSat dispersion to create efficient constellation geometries. The same deployment strategies can be used to develop a multitude of differently dispersed constellations. Different launch opportunities can be utilized to tailor a constellation for a specific region or mission objective. Constellations can also be augmented using multiple launch opportunities to optimize a constellation towards a specific mission or region. The tools developed to obtain these results can also be used to perform specific analysis on any region in order to optimize future constellations for other applications.

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.