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.

<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>

Diplomová práce se zabývá tuhostní optimalizací nosiče satelitů, tzv. Dispenseru. První kapitola uvádí přehled evropských vesmírných aktivit a poskytuje nezbytné technické pozadí týkající se nosných raket. Druhá kapitola se věnuje popisu tvorby výpočetního modelu, neboť veškeré výpočty, včetně optimalizace popsané v kapitole třetí, jsou založeny na metodě konečných prvků. Pro optimalizovanou variantu je ve čtvrté kapitole provedena základní pevnostní kontrola.

Thesis (PhD (Electrical Engineering))--Cape Peninsula University of Technology, 2019 / This dissertation presents a concept for a compact, low-profile, integrated solar panel antenna for use on small satellites in low Earth orbit. To date, the integrated solar panel antenna design approach has primarily been, patch (transparent or non-transparent) and slot radiators. The design approach presented here is proposed as an alternative to existing designs. A prototype, comprising of an optically transparent rectangular dielectric resonator was constructed and can be mounted on top of a solar panel of a Cube Satellite. The ceramic glass, LASF35 is characterised by its excellent transmittance and was used to realise an antenna which does not compete with solar panels for surface area. Currently, no closed-form solution for the resonant frequency and Q-factor of a rectangular dielectric resonator antenna exists and as a first-order solution the dielectric waveguide model was used to derive the geometrical dimensions of the dielectric resonator antenna. The result obtained with the dielectric waveguide model is compared with several numerical methods such as the method of moments, finite integration technique, radar cross-section technique, characteristic mode analysis and finally with measurements. This verification approach was taken to give insight into the resonant modes and modal behaviour of the antenna. The interaction between antenna and a triple-junction gallium arsenide solar cell is presented demonstrating a loss in solar efficiency of 15.3%. A single rectangular dielectric resonator antenna mounted on a ground plane demonstrated a gain of 4.2 dBi and 5.7 dBi with and without the solar cell respectively. A dielectric resonator antenna array with a back-to-back Yagi-Uda topology is proposed, designed and evaluated. The main beam of this array can be steered can steer its beam ensuring a constant flux density at a satellite ground station. This isoflux gain profile is formed by the envelope of the steered beams which are controlled using a single digital phase shifter. The array achieved a beam-steering limit of ±66° with a measured maximum gain of 11.4 dBi. The outcome of this research is to realise a single component with dual functionality satisfying the cost, size and weight requirements of small satellites by optimally utilising the surface area of the solar panels.

With the launch of the first CubeSat a trend of easy access to Low Earth Orbit was started. Today many educational institutes around the world design, build and operate CubeSats for educational as well as scientific purposes. This Master thesis work presents designs and development in hardware and software to achieve a flexible ground segment at the Luleå University of Technology Space Campus in Kiruna, Sweden. The existing ground station is adapted to support more frequencies and modes of operation to enable future nanosatellite projects at the university easy access to space communication. New equipment is procured and installed with existing equipment in a new location using a 19 inch rack. The thesis presents a ground segment design using software-defined radio to promote flexibility and adaptability. Software development for the ground station is carried out together with Remos Space Systems a start-up at the Arctic Business Incubator that is developing a commercial ground station software. Furthermore a brief analysis of establishing a S-band receive-only ground station at the university is conducted and a trade-off analysis regarding mission control software is made. The thesis lays the foundation and highlights future development needs for the Space Campus ground station to become operational again.

CubeSats, specially the 6U standard, is nowadays the tendency where many developers point towards. The upscaling size of the standard and payloads entail the increase of the satellite overall mass. Composite materials have demonstrated the ability to fulfill expectations like reducing structural masses, having been applied to different types of spacecraft, including small satellites. This Thesis is focused on designing, manufacturing, and dynamic testing of a 6U CubeSat made of carbon fiber, fiberglass, and aluminum. The main objective of this study was obtaining a mass reduction of a 6U CubeSat structure, maintaining the stiffness and strength. Considering the thermal effects of the used materials an outgassing test of the used materials was performed and the experimental results are presented. The CubeSat structure was entirely manufactured and tested at Cal Poly Aerospace Engineering Department facilities. A mechanical shock test and random vibration test were performed using a shock table and a shake table respectively. Results of both tests are presented. A correlation between the Experimental data and the Finite Element Model of the satellite was carried out. Finally, a comparison between 6U structure studied and aluminum 6U structures available in the market is presented.

The research presented provides an overview of a 1U+ form factor propulsion system design developed for the Cal Poly CubeSat Laboratory (CPCL). This design utilizes a Radiofrequency Electrothermal Thruster (RFET) called Pocket Rocket that can generate 9.30 m/s of delta-V with argon, and 20.2 ± 3 m/s of delta-V with xenon. Due to the demand for advanced mission capabilities in the CubeSat form factor, a need for micro-propulsion systems that can generate between 1 – 1500 m/s of delta-V are necessary. By 2019, Pocket Rocket had been developed to a Technology Readiness Level (TRL) of 5 and ground tested in a 1U CubeSat form factor that incorporated propellant storage, pressure regulation, RF power and thruster control, as well as two Pocket Rocket thrusters under vacuum, and showcased a thrust of 2.4 mN at a required 10 Wdc of power with Argon propellant. The design focused on ground testing of the thruster and did not incorporate all necessary components for operation of the thruster. Therefore in 2020, a 1U+ Propulsion Module that incorporates Pocket Rocket, the RF amplification PCB, a propellant tank, propellant regulation and delivery, as well as a DC-RF conversion with a PIB, that are all attached to a 2U customer CubeSat for a 3U+ overall form factor. This design was created to increase the TRL level of Pocket Rocket from 5 to 8 by demonstrating drag compensation in a 400 km orbit with a delta-V of 20 ± 3 m/s in the flight configuration. The 1U+ Propulsion Module design included interface and requirements definition, assembly instructions, Concept of Operations (ConOps), as well as structural and thermal analysis of the system. The 1U+ design enhances the capabilities of Pocket Rocket in a 1U+ form factor propulsion system and increases future mission capabilities as well as propulsion system heritage for the CPCL.

With the rise of CubeSats and the demonstration of their many space applications, there is interest in interplanetary CubeSats to act for example as scientific investigations or communications relays. In line with the increasing demand for this class of small satellites, the Cal Poly CubeSat Lab (CPCL) seeks to develop a bus that could support an interplanetary science payload. To facilitate this, a mission concept to conduct science of the moons of Mars, Phobos and Deimos, is investigated by determining the mission needs for a CubeSat in a Phobos-Deimos cycler orbit through the development of a baseline design to meet mission objectives. This baseline design is then compared by subsystem to CPCL’s current capabilities to identify technology, facility, and knowledge gaps and recommend a path forward to close them. The resulting baseline design is a 16U bus capable of transferring from an initial low Mars orbit to a Phobos-Deimos cycler orbit using a combined chemical and electric propulsion system. The bus is designed for a 3.5 year mission lifetime collecting radiation data and images, utilizing a relay architecture to downlink payload data. Estimates for mass, volume, and power available for an additional payload are up to 2.3 kg in ~4U with power consumption up to 13 to 38 W. This baseline requires further iteration due to non-closure of the thermal protection subsystem and improvement of other subsystems but serves as a starting point for exploration into CPCL’s next steps in becoming an interplanetary bus provider. Major subsystem areas identified for hardware performance improvement within CPCL are propulsion, communications, power, and mechanisms.

In this thesis, the design and particularities of a unique and revolution- ary scalable propulsion system are presented. A spacecraft mechanical design is included together with a mission definition, aiming to provide a context for a technology demonstration in space of an Hybrid Rocket Motor (HRM) as satellite thruster. Rocket motors have been around for many decades, with their use mainly focused on launch vehicles and large satellites, thus restricting the access to space to institutions with big budgets. To overcome this limitation, the application of a cost-effective type of rocket motor without a heritage of space utilization is explored. This is the implementation of an HRM as satellite thruster. In Chapter 2, the characteristics of this particular case of chemical rocket motor are presented in detail. The HRM applied for the present mission is a particular case of an in- house developed motor design method. As presented in Chapter 7, a scalable and versatile mechanical and propulsion design have been elab- orated following the maturation of a scalability software (Appendix A). The combination of these constitute a valuable tool allowing for a fast and accurate motor design for the desired scenario. Taking advantage of this straightforward tool, an attractive mission was defined to provide a meaningful context for the maiden use of an HRMin space. A micro satellite deep space mission, defined in Chapter 3, was chosen to validate the tool and prove Hybrid Rocket Motors (HRMs) capabilities, showing the benefits of its use over other propulsion systems already available, specifically in the small satellite family. The spacecraft design was tackled aiming to support the motor’s scalable concept while complying with the mission requirements and space standards. The out- come is an easily adaptable satellite design, justified in Chapter 8. The performed structural simulations are outlined in Appendix C to validate the developed design. Ultimately, this thesis work intends to provide the space community with a noteworthy product, opening the access to interplanetary missions provided the reduced mission costs of small satellites mounted with anHRM as propulsion system. Arising from the thesis content, research papers (Part v) have been published and presented in distinguished congresses, contributing to space development.

The unique environment of CubeSat and small satellite missions allows certain accepted paradigms of the larger satellite world to be investigated in order to trade performance for simplicity, mass, and volume. Peak Power Tracking technologies for solar arrays are generally implemented in order to meet the End-of-Life power requirements for satellite missions given radiation degradation over time. The short lifetime of the generic satellite mission removes the need to compensate for this degradation. While Peak Power Tracking implementations can give increased power by taking advantage and compensating for the temperature cycles that solar cells experience, this comes at the expense of system complexity and, given smart system design, this increased performance is negligible and possibly detrimental. This thesis investigates different Peak Power Tracking implementations and compares them to two Fixed Point implementations as well as a Direct Energy Transfer system in terms of performance and system complexity using computer simulation. This work demonstrates that, though Peak Power Tracking systems work as designed, under most circumstances Direct Energy Transfer systems should be used in small satellite applications as it gives the same or better performance with less complexity.