Body mounted germanium substrate solar cell arrays form the faces of many small satellite designs to provide the primary power source on orbit. High efficiency solar cells are made affordable for university satellite programs as triangular devices trimmed from wafer scale solar cells. The smaller cells allow array designs to pack tightly around antenna mounts and payload instruments, giving the board design flexibility. One objective of this work is to investigate the reliability of solar cells attached to FR-4 printed circuit boards. FR-4 circuit boards have significantly higher thermal expansion coefficients and lower thermal conductivities than germanium. This thermal expansion coefficient mismatch between the FR-4 board and the components causes concern for the power system in terms of failures seen by the solar cells. These failures are most likely to occur with a longer orbital lifetime and an extended exposure to harsh environments. This work compares various methods of attaching solar cells to printed circuit boards, using solder paste alone and with a silicone adhesive, and considering the application of these adhesives by comparing the solder joints when printed by screen versus a stencil. An environmental test plan was used to compare the survivability and performance of the solar arrays.

Attitude control remains one of the top engineering challenges faced by small satellite mission planning and design. Conventional methods for attitude control include propulsion, reaction wheels, magnetic torque coils, and passive stabilization mechanisms, such as permanent magnets that align with planetary magnetic fields. Drawbacks of these conventional attitude control methods for small satellites include size, power consumption, dependence on external magnetic fields, and lack of full control authority. This research investigates an alternative, novel approach to attitude-control method for small satellites, utilizing the noncommutative property of rigid body rotation sequences. Piezoelectric bimorph actuators are used to induce sinusoidal small-amplitude satellite oscillations on two of the satellites axes. While zero net change occurs on these signaled axes, the third axis can develop an average angular rate. This noncommutative attitude control methodology has several advantages over conventional methods, including scalability, power consumption, and operation outside of Earth's magnetic field. This research looks into the feasibility of such a system, and lays the foundation for a simple control system architecture.

As electronics become smaller and more capable, it has become possible to conduct meaningful and sophisticated satellite missions in a small form factor. However, the capability of small satellites and the range of possible applications are limited by the capabilities of several technologies, including attitude determination and control systems. This dissertation evaluates the use of image-based visual attitude propagation as a compliment or alternative to other attitude determination technologies that are suitable for miniature satellites. The concept lies in using miniature cameras to track image features across frames and extracting the underlying rotation. The problem of visual attitude propagation as a small satellite attitude determination system is addressed from several aspects: related work, algorithm design, hardware and performance evaluation, possible applications, and on-orbit experimentation. These areas of consideration reflect the organization of this dissertation. A “stellar gyroscope” is developed, which is a visual star-based attitude propagator that uses relative motion of stars in an imager’s field of view to infer the attitude changes. The device generates spacecraft relative attitude estimates in three degrees of freedom. Algorithms to perform the star detection, correspondence, and attitude propagation are presented. The Random Sample Consensus (RANSAC) approach is applied to the correspondence problem to successfully pair stars across frames while mitigating false-positive and false-negative star detections. This approach provides tolerance to the noise levels expected in using miniature optics and no baffling, and the noise caused by radiation dose on orbit. The hardware design and algorithms are validated using test images of the night sky. The application of the stellar gyroscope as part of a CubeSat attitude determination and control system is described. The stellar gyroscope is used to augment a MEMS gyroscope attitude propagation algorithm to minimize drift in the absence of an absolute attitude sensor. The stellar gyroscope is a technology demonstration experiment on KySat-2, a 1-Unit CubeSat being developed in Kentucky that is in line to launch with the NASA ELaNa CubeSat Launch Initiative. It has also been adopted by industry as a sensor for CubeSat Attitude Determination and Control Systems (ADCS).

This thesis presents a decentralized electromagnetic formation flying (EMFF) control method using frequency-multiplexed sinusoidal control signals. We demonstrate the EMFF control approach in open-loop and closed-loop control experiments using a single-degree-of-freedom testbed with an electromagnetic actuation system (EAS). The EAS sense the relative position and velocity between satellites and implement a frequency-multiplexed sinusoidal control signal. We use a laser-rangefinder device to capture the relative position and an ARM-based microcontroller to implement the closed-loop control algorithm. We custom-design and build the EAS that implements the formation control in one dimension. The experimental results in this thesis demonstrate the feasibility of the decentralized formation control algorithm between two satellites.

<div>As small satellites are increasingly used in the space industry, creative solutions for the use of their limited volume will be required. Conformal tanks are one idea to better make use of this volume. These tanks are non-traditionally shaped and non-axisymmetric. Because slosh can have detrimental effects on a spacecraft, it should be understood. However, slosh in these more complicated geometries has not been thoroughly investigated in the past.</div><div><br></div><div>This research looks at slosh within six geometries, five of which are conformal tanks. These geometries are evaluated in both an experiment and using CFD simulations. It was found that the total slosh motion appears to be the sum of slosh behavior along each dimension. Slosh along a line of symmetry will have center of mass movement that stays along that line. Slosh off the line of symmetry will deviate from that line unless slosh frequency is the same in each direction.</div>

The use of small satellites is rapidly growing, especially satellites with masses between 1 and 10 kg and few litres of volume. The main reasons are due to the low development time and cost. Electronics miniaturization and high density integration is enabling the small satellites class to perform more and better tasks and at a lower cost. When deployable structures are required for the missions, the actual paradigm is that there are very few that have been successfully developed and flown. It is usually not possible to scale down existing deployable structures from larger satellites. Power and attitude control is also very limited in small satellites thus, completely new deployable structures, low mass and with high packaging ratio (yet large and with adequate mechanical properties when deployed) must be developed. Furthermore, such new structures are usually made of very thin and light materials which complicates the on-ground tests prior the launch. Therefore, advances in modelling and simulation deployable structures such as booms are also of great interest for the scientific community. This thesis and the papers included herein focus on the finite element modelling of a meter-class passively deployable boom – based on the SIMPLE boom by Thomas W. Murphey – and deployment simulations. Experimental tests were also carried on a boom prototype suspended from a gravity off-loading system. An analytical model produced certain parameters which are used for validation of the finite element model. The strain energy stored in the boom prior to deployment and spacecraft displacements during deployment agreed well. The deployment time, however, have discrepancies: the models predicted a deployment time six times faster than the experimental tests. For that reason the deployment simulations cannot be compared with the tests. The reason of the discrepancies are believed to be due to the actual material model and the contacts formulation used in the finite element model. The finite element simulations, however, shows a reasonable behaviour given the nature of the deployment thus, despite the necessary improvements, we believe that future improvements in the material and friction models will provide us more realistic results. / Användningen av små satelliter ökar snabbt, särskilt satelliter med en vikt på mellan 1 och 10 kg och bara några liters volym. De främsta orsakerna till detta är den korta utvecklingstiden och den låga kostnaden. Elektronikminiatyrisering och hög integreringsdensitet möjliggör för små satelliter att utföra fler och bättre uppgifter till en lägre kostnad. När utfällbara strukturer krävs för uppdragen är nuvarande läge att det är få som utvecklats och flugits framgångsrikt. Det är inte heller alltid möjligt att skala ner utfällbara strukturer som utformats för användning i större satelliter. I små satelliter är den tillgängliga elektriska energin och volymen starkt begränsade faktorer och därmed måste helt nya passivt utfällbara strukturer med låg vikt och liten packningsvolym, men ändå rätt storlek och mekaniska egenskaper när de är utfällda, utvecklas. Dessa strukturer är vanligen tillverkade av mycket tunna och lätta material, som komplicerar tester innan uppskjutningen p.g.a. tyngdkraften. Därför är det av stort intresse att noggrant kunna modellera och simulera ett tyngdlöst utfällningsförlopp. Denna licentiatuppsats och bilagda artiklar i fokuserar på finit elementmodellering och utfällningssimuleringar av en 1 meter lång passivt utfällbar bom baserad på SIMPLE-bommen som utformats av Thomas W. Murphey. Utfällningsexperiment har utförts på en prototyp av bommen upphängd i ett tyngdkraftskompenserande system. Analytiska modeller har använts för att validera simuleringarna och töjningsenergin som lagrats i bommen innan utfällning och rymdfarkostens förflyttning efter utfällning överensstämmer väl. Utfällningstiden avviker dock och båda modellerna predikterar en utfällningstid som är sex gånger snabbare än den tiden som observeras i experimenten. Anledningen till skillnaderna antas delvis bero på begränsningar i den använda materialmodellen och i algoritmer för hantering av kontakt i den finita elementmodellen. De finite elementsimuleringarna visar dock ett rimligt dynamisk beteende hos bommen baserat på vad som observerats i experimenten och även om modellen är i behov av förbättring så finns det stora förhoppningar att åstadkomma en mer realistisk modell genom införande av förbättrade kontakalgoritmer och nogrannare modellering av dämpning och friktion. / <p>QC 20130506</p>

<p>  </p> <p>As space agencies look to conduct more scientific missions beyond Earth orbit, low-cost access to space becomes indispensable. Small satellites (smallsats) fulfill this need as they can be developed at a fraction of the cost of traditional large satellites. Consequently, smallsats are being envisioned for planetary science missions at several destinations including Mars. However, a significant challenge for interplanetary smallsats is performing fully-propulsive orbit insertion because modern smallsat propulsion technologies have limited total velocity change capabilities. At destinations with significant atmospheres, this challenge can be circumvented via <em>aerocapture</em>, a technique that uses a single atmospheric pass to convert a hyperbolic approach trajectory into a captured elliptical orbit. Aerocapture has been shown to enable significant propellant mass savings as compared to fully-propulsive orbit insertion, making it an attractive choice for smallsats. Performing aerocapture with smallsats requires a suitable vehicle design that satisfies the associated control requirements and volumetric constraints. To address this requirement, this dissertation proposes the <em>morphable entry system </em>(MES), a conceptual deployable entry vehicle that utilizes shape morphing to follow a desired atmospheric flight profile during aerocapture. The aerocapture performance of the MES at Mars is investigated using a six degree-of-freedom aerocapture simulation environment. The shape morphing strategy employed by the MES is shown to be feasible for targeting desired angle of attack and sideslip angle profiles that lead to successful orbit captures. Furthermore, the robustness of the MES to simulated day-of-flight uncertainties while employing angle of attack control is demonstrated through a Monte Carlo dispersion analysis. The major contributions of this research as well as areas of future work are described.</p>