Cubesats are picosatellites that have a mass of less than 1.3kg and have a shape of acube. As a result of their low cost of development and launch, cubesats are gainingpopularity in industry and academia. These satellites are also a cost-efective way forspace technology demonstrations. University of Würzburg has a longstanding cubesatprogram started with the launch of UWE-1 in 2005. This was followed by UWE-2 andUWE-3. Several technologies were tested and validated using the UWE platform. Thelast mission UWE-3 has successfully tested an attitude control system.In the next mission, UWE-4 will demonstrate an orbit control system. Being a picosatellite as small as this one (10 x 10 x 10cm 3 and 1kg) brings new challenges intodi↵erent aspects of satellite design, development, control and operation. The orbit con-trol of such a satellite is one of the problems that should be tackled. Being such a smallsatellite means having less propellant mass and much smaller thrusters than conventionalsatellites. These should be addressed in the orbit control. UWE-4 will take advantage of four NanoFEEP thrusters, on one side. Because of theiraccuracy and functionality, these thrusters can be used to implement a continuous thrustsystem. They are also a good choice because of their low energy usage. This work startswith the preparation that was needed to implement a control system. Then explains thestate of the art for continuous thrust control systems. Implements two di↵erent methods,based on perfect control and discusses the outcome. It discuses the limiting factors, likefuel mass, available electrical energy and their e↵ect on the controller performance andconcludes with recommendation for the future researches. / UWE-4

The development of a small spacecraft attitude determination and control subsystem is described. This subsystem is part of The Space Flight Laboratory's Generic Nanosatellite Bus. With a 20cm3 body, the bus has an attitude determination and control subsystem capable of full three-axis stabilization and control enabling more advanced missions previously only possible with bulkier and more power-consuming attitude control hardware. Specific contributions to the Space Flight Lab's attitude control hardware are emphasised. Particularly, the full development of a 32g three-axis nanosatellite rate sensing unit is described. This includes embedded software development, skew calibration, hardware modeling and qualification testing for the unit. Development work on a three-axis boom-mounted magnetometer is also detailed. A full hardware design is also described for a new microsatellite-sized rate sensor. Larger and more powerful than the nanosatellite rate sensors, the design ensures a low noise, low drift architecture to improve attitude determination on future microsatellite missions.

The development of a small spacecraft attitude determination and control subsystem is described. This subsystem is part of The Space Flight Laboratory's Generic Nanosatellite Bus. With a 20cm3 body, the bus has an attitude determination and control subsystem capable of full three-axis stabilization and control enabling more advanced missions previously only possible with bulkier and more power-consuming attitude control hardware. Specific contributions to the Space Flight Lab's attitude control hardware are emphasised. Particularly, the full development of a 32g three-axis nanosatellite rate sensing unit is described. This includes embedded software development, skew calibration, hardware modeling and qualification testing for the unit. Development work on a three-axis boom-mounted magnetometer is also detailed. A full hardware design is also described for a new microsatellite-sized rate sensor. Larger and more powerful than the nanosatellite rate sensors, the design ensures a low noise, low drift architecture to improve attitude determination on future microsatellite missions.

Within the first decade of the 21st century, the aerospace community has seen many more opportunities to launch small spacecraft in the 10 to 100 kg mass class. Coupled with this has been consistent interest from the government in developing small-spacecraft platforms to expand civil and military mission possibilities. Small spacecraft have also given small organizations such as universities an increased access to space. Because small satellites are limited in size, power, and mass, new and often nontraditional capabilities must be explored and developed to make them viable and attractive when compared with larger and more proven spacecraft. Moreover, small organizations that wish to contribute technically are often limited by the small size of their teams and available resources, and need creative solutions for meeting mission requirements. A key need is in space-to-ground communications. Complex missions typically require large amounts of data transfer to the ground and in a timely fashion. Available options trade hardware cost, available ground stations or networks, available operating-frequency range, data-rate performance, and ease of use. A system for small spacecraft will be presented based upon Radio Frequency Integrated Circuits (RFIC) that minimizes development effort and maximizes interface control to meet typical small-spacecraft communications requirements. RFICs are low-cost components that feature pre-built radio hardware on a chip that can be expanded easily by developers with little or no radio experience. These devices are widespread in domestic applications for short-range connectivity. A preliminary design and prototype is presented that meets basic spaceflight requirements, offers data rates in the 55 to 85 kbps range, and has completed basic proof-of-concept testing. While there are higher-data-rate alternatives in existence, the solution presented here strikes a useful balance among data rate, parts cost, and ease of use for non experts, and gives the user operational control necessary to make air-to-ground communications time effective.

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.

O uso de pequenos satélites tem aumentado substancialmente nos últimos anos devido ao custo reduzido de desenvolvimento e lançamento, assim como pela flexibilidade oferecida pela utilização de componentes comerciais. Este trabalho propõe o projeto e a implementação de uma plataforma para teste, controle e qualificação de circuitos integrados (Integrated Circuits, CIs) comerciais e customizados para uso em aplicações espaciais. Esta plataforma flexível pode ser ajustada a uma gama de dispositivos e interfaces, e reduz os esforços de integração desses componentes e, portanto, acelera o desenvolvimento de todo o projeto. O sistema proposto é sintetizado em um tecnologia de Arranjo de Portas Programáveis em Campo (Field Programmable Gate Array) baseado em memória Flash, que, apesar de não ser classificado para uso aeroespacial, testes demonstram a viabilidade de seu uso. Este sistema adaptável permite o controle de novas cargas-úteis e softcores para o teste e validação antes da sua aplicação em voo. A comunicação com dispositivos é feita através de protocolos préimplementados. Os resultados de testes funcionais in loco sugerem a possibilidade de aplicação desta plataforma para uso em Cubesats. A primeira aplicação desta plataforma foi no teste do controle da placa de carga-útil do NanoSatC-BR1, o primeiro nanossatélite científico brasileiro, lançado em órbita em 2014. / The number of small satellites has substantially increased in the last years due to reduced development and launching costs, as well as due to the flexibility brought by the usage of commercial off the shelf components. This work purposes the design and implementation of a platform for test, control and qualification of commercial and customized integrated circuits for space applications. This flexible platform can be adjusted to control a wide range of devices and interfaces, and is intended to reduce the integration difficulties, resulting in the speed up of some of the project stages. The platform is synthesized in a Flash-based Field Programmable Gate Array technology. The target device is not qualified for aerospace projects. Nevertheless, previous radiation tests demonstrated its hardness for space missions. The system is adaptable and makes it possible to control, test and validate new payloads and softcores before flight. The communication between devices is done through pre-implemented protocols. Functional tests suggested the possibility to apply the platform in Cubesats projects. The first application of this platform was in the NanoSatC-BR1, the first Brazilian scientific nanosatellite, to test the controller of the payload board.

Differential drag controllers have been examined in the context of a full system simulation of a target/chaser pair of spacecraft in low Earth orbit. An Extended Kalman Filter has been designed to process measurement sets from GPS receivers on the target and chaser spacecraft. The estimated state from the Kalman Filter is used in a differential drag control algorithm to determine the appropriate control action. Modifications are made to the standard differential drag control algorithms to reduce unnecessary actuations in the presence of errors in the dynamic modeling, control actuation, and incoming measurements. Detailed explanations of the algorithms, dynamic models, and derivations for both the Kalman Filter and the differential drag control laws are presented. Multiple test cases are used to validate the controller performance under a variety of initial conditions. In these simulations, the differential drag control algorithms successfully maneuver the chaser spacecraft from the initial conditions to a final state with instantaneous time-average position (relative to the target spacecraft) of not more than 10 meters in the radial and in-track directions. Modifications to the standard control algorithms ensure that extraneous control actuations are minimized. An optimization algorithm is used determine the time-optimal differential drag control history, and the results are compared to the standard control logic and modified control logic. Based on the optimization results, it is recommended that a system employing differential drag control (especially those with limited computational resources) should use the modified control logic, as it provides a standardized methodology that can be followed in any mission. / text

The attitude determination and control subsystem (ADCS) is a critical part of any satellite conducting scientific experiments that require accurate positioning (such as Earth observation and solar spectroscopy). The engineering design process of this subsystem has a long heritage; yet, it is surrounded by several limitations due to the stringent physical constraints imposed on small satellites. These limitations (e.g., limited computational capabilities, power, and volume) require an improved approach for the purpose of attitude determination (AD) and control. Previous space missions relied mostly on the extended Kalman filter (EKF) to estimate the relative orientation of the spacecraft because it yields an optimal estimator under the assumption that the measurement and process models are white Gaussian processes. However, this filter suffers from several limitations such as a high computational cost. This thesis addresses all the limitations found in small satellites by introducing a computationally efficient algorithm for AD based on a fuzzy inference system with a gradient decent optimization technique to calculate and optimize the bounds of the membership functions. Also, an optimal controller based on a fractional proportional-integral-derivative controller has been implemented to provide an energy-efficient control scheme. The AD algorithm presented in this thesis relies on the residual information of the Earth magnetic field. In contrast to current approaches, the new algorithm is immune to several limitations such as sensitivity to initial conditions and divergence problems. Additionally, its computational cost has been reduced. Simulation results illustrate a higher pointing stability, while maintaining satisfying levels of pointing accuracy and increasing reliability. Moreover, the optimal controller designed provides a shorter time delay, settling time, and steady-state error. This demonstrates that accurate attitude determination and control can be conducted in small spacecraft.

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.

O uso de pequenos satélites tem aumentado substancialmente nos últimos anos devido ao custo reduzido de desenvolvimento e lançamento, assim como pela flexibilidade oferecida pela utilização de componentes comerciais. Este trabalho propõe o projeto e a implementação de uma plataforma para teste, controle e qualificação de circuitos integrados (Integrated Circuits, CIs) comerciais e customizados para uso em aplicações espaciais. Esta plataforma flexível pode ser ajustada a uma gama de dispositivos e interfaces, e reduz os esforços de integração desses componentes e, portanto, acelera o desenvolvimento de todo o projeto. O sistema proposto é sintetizado em um tecnologia de Arranjo de Portas Programáveis em Campo (Field Programmable Gate Array) baseado em memória Flash, que, apesar de não ser classificado para uso aeroespacial, testes demonstram a viabilidade de seu uso. Este sistema adaptável permite o controle de novas cargas-úteis e softcores para o teste e validação antes da sua aplicação em voo. A comunicação com dispositivos é feita através de protocolos préimplementados. Os resultados de testes funcionais in loco sugerem a possibilidade de aplicação desta plataforma para uso em Cubesats. A primeira aplicação desta plataforma foi no teste do controle da placa de carga-útil do NanoSatC-BR1, o primeiro nanossatélite científico brasileiro, lançado em órbita em 2014. / The number of small satellites has substantially increased in the last years due to reduced development and launching costs, as well as due to the flexibility brought by the usage of commercial off the shelf components. This work purposes the design and implementation of a platform for test, control and qualification of commercial and customized integrated circuits for space applications. This flexible platform can be adjusted to control a wide range of devices and interfaces, and is intended to reduce the integration difficulties, resulting in the speed up of some of the project stages. The platform is synthesized in a Flash-based Field Programmable Gate Array technology. The target device is not qualified for aerospace projects. Nevertheless, previous radiation tests demonstrated its hardness for space missions. The system is adaptable and makes it possible to control, test and validate new payloads and softcores before flight. The communication between devices is done through pre-implemented protocols. Functional tests suggested the possibility to apply the platform in Cubesats projects. The first application of this platform was in the NanoSatC-BR1, the first Brazilian scientific nanosatellite, to test the controller of the payload board.