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This work aims to implement 3D microstructures that generate light with orbital angular momentum towards applications in Biophotonics. Over the past few decades, 3D printing has established itself as the most versatile technology with effortless adaptability. Parallel to this, the concept of miniaturiza tion has seen tremendous growth irrespective of the field and has become an estab lished trend motivated by the need for compact, portable and multi-function devices. Therefore, when these two concepts get together, i.e., 3D printing of miniaturized objects, it could lead to an exciting path with endless opportunities. When it comes to optics, miniaturized 3D printing offers the potential to create compact optical micro-systems and exhibits a way to manufacture freeform µ-optics. In particular, two-photon lithography (TPL) is a cutting edge 3D printing technology that has re cently demonstrated groundbreaking solutions for optics as it offers high resolution with a great degree of flexibility. With a TPL 3D printer, it is possible to fabricate complex µ-optical elements and employ them for compelling applications. In recent years, light with orbital angular momentum (OAM), or ”twisted” light, has captured the interests of several researchers due to its inspiring applications. Tra ditionally, to generate OAM beams, one would require bulk, table-top optics, restrict ing their applications to over-the-table setup. An alternative approach of OAM beam generation is through µ-structures over the fiber, as they can open up new opportu nities, especially in Bioscience, and facilitate in-vivo operations. In particular, this probe-like setup can be used for processes such as optical trapping, high-resolution microscopy, etc. Hence, I propose the development of a novel approach with un precedented capabilities for generating OAM beams right from single-mode optical fibers, by transforming its Gaussian-like output beam by using complex 3D printed microstructures. In this document, I will showcase designs and results on generating Bessel beams (both zeroth- and high-order) and high-NA converging beams (with and without OAM) for optical trapping from the fiber. Remarkably, I achieved the first-ever fiber-based high-order Bessel beam generation and the first-ever fiber optical tweezers with OAM.

Recent interest surrounding large scale satellite constellations has increased analysis efforts to create the most efficient designs. Multiple studies have successfully optimized constellation patterns using equations of motion propagation methods and genetic algorithms to arrive at optimal solutions. However, these approaches are computationally expensive for large scale constellations, making them impractical for quick iterative design analysis. Therefore, a minimalist algorithm and efficient computational method could be used to improve solution times. This thesis will provide a tool for single target constellation optimization using spherical trigonometry propagation, and an evolutionary genetic algorithm based on a multi-objective optimization function. Each constellation will be evaluated on a normalized fitness scale to determine optimization. The performance objective functions are based on average coverage time, average revisits, and a minimized number of satellites. To adhere to a wider audience, this design tool was written using traditional Matlab, and does not require any additional toolboxes. To create an efficient design tool, spherical trigonometry propagation will be utilized to evaluate constellations for both coverage time and revisits over a single target. This approach was chosen to avoid solving complex ordinary differential equations for each satellite over a long period of time. By converting the satellite and planetary target into vectors of latitude and longitude in a common celestial sphere (i.e. ECI), the angle can be calculated between each set of vectors in three-dimensional space. A comparison of angle against a maximum view angle, , controlled by the elevation angle of the target and the satellite’s altitude, will determine coverage time and number of revisits during a single orbital period. Traditional constellations are defined by an altitude (a), inclination (I), and Walker Delta Pattern notation: T/P/F. Where T represents the number of satellites, P is the number of orbital planes, and F indirectly defines the number of adjacent planes with satellite offsets. Assuming circular orbits, these five parameters outline any possible constellation design. The optimization algorithm will use these parameters as evolutionary traits to iterate through the solutions space. This process will pass down the best traits from one generation to the next, slowly evolving and converging the population towards an optimal solution. Utilizing tournament style selection, multi-parent recombination, and mutation techniques, each generation of children will improve on the last by evaluating the three performance objectives listed. The evolutionary algorithm will iterate through 100 generations (G) with a population (n) of 100. The results of this study explore optimal constellation designs for seven targets evenly spaced from 0° to 90° latitude on Earth, Mars and Jupiter. Each test case reports the top ten constellations found based on optimal fitness. Scatterplots of the constellation design solution space and the multi-objective fitness function breakdown are provided to showcase convergence of the evolutionary genetic algorithm. The results highlight the ratio between constellation altitude and planetary radius as the most influential aspects for achieving optimal constellations due to the increased field of view ratio achievable on smaller planetary bodies. The multi-objective fitness function however, influences constellation design the most because it is the main optimization driver. All future constellation optimization problems should critically determine the best multi-objective fitness function needed for a specific study or mission.

<p>Increasingly, space flight missions are planned to traverse regions of space with complex dynamical environments influenced by multiple gravitational bodies. The nature of these systems produces motion and regions of sensitivity that are, at times, unintuitive,</p> <p>and the accumulation of trajectory dispersions from a variety of sources guarantees that spacecraft will deviate from their pre-planned trajectories in this complex environment, necessitating the use of a targeting process to generate a new feasible reference path. To ensure mission success and a robust path planning process, trajectory designers require insight into the interaction between the targeting process, the baseline trajectory, and the dynamical environment. In this investigation, the convergence behavior of these targeting processes is examined. This work summarizes a framework for characterizing and predicting the convergence behavior of perturbed targeting problems, consisting of a set of constraints, design variables, perturbation variables, and a reference solution within a dynamical system. First, this work identifies the typical features of a convergence basin and identifies a measure of worst-case performance. In the absence of an analytical method, efficient numerical discretization procedures are proposed based on the evaluation of partial derivatives at the reference solution to the perturbed targeting problem. A method is also proposed for approximating the tradespace of position and velocity perturbations that achieve reliable</p> <p>convergence toward the baseline solution. Additionally, evaluated scalar quantities are introduced to serve as predictors of the simulation-measured worst-case convergence behavior based on the local rate of growth in the constraints as well as the local relative change in the targeting-employed partial derivatives with respect to perturbations.</p> <p><br></p> <p>A variety of applications in different dynamical regions and force models are introduced to evaluate the improved discretization techniques and their correlation to the predictive metrics of convergence behavior. Segments of periodic orbits and transfer trajectories from past and planned missions are employed to evaluate the relative convergence performance across sets of candidate solutions. In the circular restricted three-body problem (CRTBP), perturbed targeting problems are formulated along a distant retrograde orbit and a near-rectilinear halo orbit (NRHO) in the Earth-Moon system. To investigate the persistence of results from the CRTBP in an ephemeris force model, a targeting problem applied to an NRHO is analyzed in both force models. Next, an L1 -to-L2 transit trajectory in the Sun-Earth system is studied to explore the effect of moving a maneuver downstream along</p> <p>a trajectory and altering the orientations of the gravitational bodies. Finally, a trans-lunar return trajectory is explored, and the convergence behavior is analyzed as the final maneuver time is varied.</p>

As an alternative solution to global warming, this thesis explores the possibility of aspace-based geoengineering scheme that may prove worthwhile to implement in parallel toother environmental efforts that help mitigate impact of climate change. One suggestionof a geoengineering solution is deploying a large number of sunshades in the vicinity ofthe first Lagrange point of the Sun-Earth system, and this prospective sunshade projectwould serve to shield Earth from incident solar radiation. This thesis is an extension ofa feasibility study for the implementation of this large-scale mission, and has a focus oncomparing electric thrusters to solar sailing as a means of propulsion. Background onelectric propulsion systems and spaceflight mechanics is provided. The investigation wasperformed by defining the spacecraft configurations, and then computing trajectories toa point of escape from Earth and from there to the final equilibrium point.Our results show that in order to meet the propellant demands of the electric thrusters,the launch mass would need to increase by around 15-25 % compared to the solar sailingimplementation, equating to around 1010 kg. Nevertheless, electric propulsion could stillbe a beneficial choice since it would allow shorter transfer times for each shade whichreduces the radiation exposure and subsequent degradation of the spacecraft’s systems.It was found that the transfer time with electric propulsion would be about one-half orone-fifth that of solar sailing, depending on spacecraft parameters. Additionally, electricpropulsion allows a much lower initial parking orbit, and while this would increase the ra-diation exposure it would also reduce the launch costs due to the higher payload capacityto lower altitudes. However, electric propulsion of this scale require prior advancementsin xenon or other inert propellant extraction methods and possibly a wide-scale construc-tion of air separation plants.

This document describes the design and analysis of the Navigation, Guidance and Control System for the KnightSat project. The purpose for the project is to test and demonstrate new technologies the Air Force would be interested in for research and development. The primary mission of KnightSat is to show how a constellation of satellites can maintain relative position with each other autonomously using the Microwave Electro Thermal (MET) thruster. The secondary mission is to use multiple satellite imagery to obtain 3 dimensional stereo photographs of observable terrain. Formation flying itself has many possible uses for future applications. Selected missions that require imaging or data collection can be more economically accomplished using smaller multiple satellites. The MET thruster is a very efficient, but low thrust alternative that can provide thrust for a very long time, hence provide the low thrust necessary to maintain the satellites at a constant separation. The challenge is to design a working control algorithm to provide the desired output data to be used to command the MET thrusters. The satellites are to maintain a constant relative distance from each other, and use the least amount of fuel possible. If one satellite runs out of fuel before the other, it would render the constellation less useful or useless. Hence, the satellites must use the same amount of fuel in order to maintain an optimal operational duration on orbit.

Once a satellite has completed its operational period, it must be removed responsibly in order to reduce the risk of impacting other missions. Geostationary Transfer Orbits (GTOs) offer unique challenges when considering disposal of spacecraft, as high eccentricity and orbital energy give rise to unique challenges for spacecraft designers. By leveraging small satellite research and integration techniques, a deployable drag sail module was analyzed that can shorten the expected orbit time of launch vehicle stages in GTO. A tool was developed to efficiently model spacecraft trajectories over long periods of time, which allowed for analysis of an object’s expected lifetime after its operational period had concluded. Material limitations on drag sail sizing and performance were also analyzed in order to conclude whether or not a system with the required orbital performance is feasible. It was determined that the sail materials and configuration is capable of surviving the expected GTO environment, and that a 49 m2 drag sail is capable of sufficiently shortening the amount of time that the space vehicles will remain in space.

There is an increased need for spacecraft to quickly and efficiently de-orbit themselves as the amount of debris in orbit around Earth grows. Defunct spacecraft pose a significant threat to the LEO environment due to their risk of fragmentation. If these spacecraft are de-orbited at the end of their useful life their risk to future spacecraft is greatly lessened. A proposed method of efficiently de-orbiting spacecraft is to use an inflatable thin-film envelope to increase the body's area to mass ratio and thusly shortening its orbital lifetime. The system and analysis presented in this project is sized for use on a CubeSat as they are an effective utility as a technology demonstration platform. Analysis has been performed to characterize the orbital dynamics of high area to mass ratio spacecraft as well as the leak rate of such an inflatable device in a vacuum environment. Results show that a 1U CubeSat can be de-orbited using a 1.7 meter diameter spherical device in just under one year while using 0.7 grams of inflating gas, this is compared to over 25 years without any method of post-mission disposal.