Two immiscible uids converging at microchannel cross-junction results in the for- mation of periodic, dispersed microslugs. This microslug formation phenomenon has been proposed as the basis for a fuel injection system in a novel, discrete mono- propellant microthruster design for use in next-generation nanosatellites. Previous experimental work has demonstrated the ability to repeatably generate fuel slugs with characteristics commensurate with the intended application. In this work, numerical modeling and simulation are used to further study this problem, and identify the sensitivity of the slug characteristics to key material properties including surface ten- sion, contact angle and fuel viscosity. These concerns are of practical concern for this application due to the potential for thermal variations and/or uid contamination during typical operation. For each of these properties, regions exist where the slug characteristics are essentially insensitive to property variations. Future microthruster system designs should target and incorporate these stable ow regions in their baseline operating conditions to maximize robustness of operation.

Experimental investigations have evaluated the feasibility of using laser-driven plasma microthrusters for small-thrust, high-specific-impulse space maneuvers, particularly for micro- and nanosatellite missions. Recent work made use of the Mach2 hydromagnetics code for the construction of an adequate computational model of the micro-thruster opera- tion. This thesis expounds on this previous work by extending the computational modeling capabilities, allowing for the determination of plasma plume properties and characteristic performance assessment of the microthruster; this allows for further computational investi- gation of the performance improvements achieved by new design considerations. Two par- ticular design changes are implemented and measured: (i) the simulation of microthruster performance intentionally achieving laser-supported detonation of energetic polymer fuels for higher-thrust capabilities, and (ii) the implementation of an axisymmetric nozzle to improve passive solid-fuel performance. The Mach2 hydromagnetics code with the new performance assessment capabilities was used to examine the performance improvement of these new modes of operation; results of the simulations are presented and then evaluated for their use in the overall design of the plasma microthruster. Laser-supported detona- tion shows a tremendous potential increase in the laser momentum coupling coefficient Cm , and demonstrates a much higher thrust; the axisymmetric nozzle varies with nozzle half-angle and length, but still demonstrates expected nozzle trends and improves the laser momentum coupling coefficient, Cm , by up to 230% for some designs considered.

Experimental investigations have evaluated the feasibility of using laser-driven plasma microthrusters for small-thrust, high-specific-impulse space maneuvers, particularly for micro- and nanosatellite missions. Recent work made use of the Mach2 hydromagnetics code for the construction of an adequate computational model of the micro-thruster opera- tion. This thesis expounds on this previous work by extending the computational modeling capabilities, allowing for the determination of plasma plume properties and characteristic performance assessment of the microthruster; this allows for further computational investi- gation of the performance improvements achieved by new design considerations. Two par- ticular design changes are implemented and measured: (i) the simulation of microthruster performance intentionally achieving laser-supported detonation of energetic polymer fuels for higher-thrust capabilities, and (ii) the implementation of an axisymmetric nozzle to improve passive solid-fuel performance. The Mach2 hydromagnetics code with the new performance assessment capabilities was used to examine the performance improvement of these new modes of operation; results of the simulations are presented and then evaluated for their use in the overall design of the plasma microthruster. Laser-supported detona- tion shows a tremendous potential increase in the laser momentum coupling coefficient Cm , and demonstrates a much higher thrust; the axisymmetric nozzle varies with nozzle half-angle and length, but still demonstrates expected nozzle trends and improves the laser momentum coupling coefficient, Cm , by up to 230% for some designs considered.

The need for micropropulsion solutions for spacecraft has been steadily increasing as scientific payloads require higher accuracy maneuvers and as the use of small form-factor spacecraft such as CubeSats becomes more common. Of the technologies used for this purpose, electrospray thrusters offer performance that make them an ideal choice. Electrosprays offer high accuracy impulse bits at low power and high efficiency, and have low volume requirements. Design choice reasoning and preliminary testing results are presented for two electrospray thruster designs. The first thruster, named the Demonstration thruster, is operated in atmospheric conditions and serves as a highly visible example of the basic concepts of electrospray technology applied to micropropulsion. It features a single capillary needle emitter and the acetone propellant flow is driven actively by a syringe pump. The second thruster, named the Research thruster, is operated in the vacuum environment and is designed for modularity for its expected use in future research efforts. Propellant flow is also driven actively using a syringe pump. Initial configuration of the Research thruster is a linear array of five capillary needle emitters, though testing is conducted with only one emitter in this thesis. Tests using un-doped glycerol and sodium iodide doped glycerol (20% by weight) are conducted for the Research thruster. Both thruster designs use stainless steel 18 gauge blunt dispensing needles (0.038 in / 0.965 mm ID) as their emitters. Applied voltage to the emitter(s) relative to the grounded extractor is swept from 2100 V to 3700 V for the Demonstration thruster testing and from 4000 V to 4500 V for the Research thruster. Currents incident on a collection plate downstream of the emission plume and on the extractors of the thrusters were measured directly with a pico-ammeter. Measurements made during testing of the Demonstration thruster are inconsistent due to charge loss as propellant travels through the air, though currents as high as 5.1x10-9 A on the collection plate and 2x10-7 A on the extractor are recorded. Currents for Research thruster testing using un-doped glycerol were measured as high as 4.9x10-8 A on the collection plate and 5x10-9 A on the extractor, showing an interception rate as high as 17%. Currents using sodium iodide doped glycerol were measured as high as 7x10-7 A on the collection plate. Discussion is given for the visual qualities of cone-jet emission for all testing. Keywords:

Propulsive capabilities on a CubeSat are the next step in advancement in the Aerospace Industry. This is no longer a quest that is being sought by just university programs, but a challenge that is being taken on by all of the industry due to the low-cost missions that can be accomplished. At this time, all of the proposed micro-thruster systems still require some form of development or testing before being flight-ready. Stellar Exploration, Inc. is developing a monopropellant micropropulsion system designed specifically for CubeSat application. The addition of a thruster to a CubeSat would expand the possibilities of what CubeSat missions are capable of achieving. The development of these miniature systems comes with many challenges. One of the largest challenges that a hot thruster faces is the ability to complete burns for the specified mission without transferring excessive heat into the propulsion tank. Due to the close proximity of the thruster to the tank, thermal standoff options are necessary to help alleviate the heat going through the system, especially while in a thermally extreme environment. This thesis examines the heat transfer that occurs within a CubeSat with an operating hydrazine monopropellant thruster. Thermal analysis of the system revealed that having a solid stainless steel barrier between the thruster and tank led to increasing temperatures greater than 400K in the propellant tank while in an environment exposed to the sun. This creates a large amount of risk for the CubeSat and its mission. The use of a thermal insulating material or a hollow barrier for the standoff decreased the risk of using this system. This creates a standoff where the heat of the propellant reaction does not reach the propellant in the tank. Therefore, the maximum temperature that the tank reaches is equivalent to the temperature of the external environment while in extreme conditions. These results create the confidence that the thermal standoffs will function as intended to protect the spacecraft and its payload during flight.

Supersonic micronozzles operate in the unique viscosupersonic flow regime, characterized by large Mach numbers (M>1) and low Reynolds numbers (Re<1000). Past research has primarily focused on the design and analysis of converging-diverging de Laval nozzles; however, plug (i.e. centerbody) designs also have some promising characteristics that might make them amenable to microscale operation. In this study, the effects of plug geometry on plug micronozzle performance are examined for the Reynolds number range Re = 80-640 using 2D Navier-Stokes-based simulations. Nozzle plugs are shortened to reduce viscous losses via three techniques: one - truncation, two - the use of parabolic contours, and three - a geometric process involving scaling. Shortened nozzle are derived from a full length geometry designed for optimal isentropic performance. Expansion ratio (ε = 3.19 and 6.22) and shortened plug length (%L = 10-100%) are varied for the full Reynolds number range. The performance of plug nozzles is then compared to that of linear-walled nozzles for equal pressure ratios, Reynolds numbers, and expansion ratios. Linear-walled nozzle half-angle is optimized to to ensure plug nozzles are compared against the best-case linear-walled design. Results indicate that the full length plug nozzle delivers poor performance on the microscale, incurring excessive viscous losses. Plug performance is increased by shortening the nozzle plug, with the scaling technique providing the best performance. The benefit derived from reducing plug length depends upon the Reynolds number, with a 1-2% increase for high Reynolds numbers an up to 14% increase at the lowest Reynolds number examined. In comparison to Linear-walled nozzle, plug nozzles deliver superior performance when under-expanded, however, this trend reverses at low pressure ratios when the nozzles become over-expanded.

<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>

<p>This thesis presents work on two major areas of research. The first area of research involves the use of a dual-nozzle piezoelectric inkjet printing system to print bilayer aluminum bismuth (III) oxide nanothermite samples. The combinatorial printing method allows for separate fuel and oxidizer inks to be printed adjacent to each other at prescribed offset distances. The effect of the bilayer thickness on the burning rate of the samples is investigated using high-speed imaging. Analysis of the burning rate data revealed that there is no statistically significant relationship between these two parameters. This result was used to determine the dominant processes that control the propagation rate in nanothermite systems. It was concluded that convective processes dominate the burning rate rather than diffusive processes. The second area of research involved synthesizing inks suitable for inkjet printing using two promising high nitrogen energetic materials called BTATz and DAATO_3.5. The performance of the developed inks was characterized using four experiments. The thermal stability and exothermic behavior of the inks were determined using DSC and TGA analysis. The results revealed that the inks are more thermally stable than the base materials. The inks were used to print lines that were subsequently used to determine burning rates. DAATO_3.5 samples were determined to have faster burning rates than BTATz. Closed pressure bomb experiments were conducted to determine the gas producing capability of the high nitrogen inks. BTATz samples showed better performance in terms of peak static pressures and pressurization rates. 3D printed microthrusters were developed to test the thrust performance of the inks. Peak thrust, total impulse, and specific impulse values are reported and were determined to be suitable for use with Class 1 micro-spacecraft. Finally, a microthruster array prototype was developed to demonstrate the capability to use additive manufacturing to create high packing density arrays.</p>

Les travaux de cette thèse ont porté sur la conception, la réalisation et la caractérisation de matrices de micro-propulseurs à propergol solide intégrés sur silicium. Ces structures sont dédiées la stabilisation de drone miniature et pouvant aussi être utilisées pour la propulsion des Micro/Nano-Satellites. Les travaux se sont effectués dans le cadre d'un projet financé par la Direction Générale pour l'Armement (DGA) en collaboration entre le LAAS-CNRS et la société PROTAC du groupe THALES. Le principe de fonctionnement d'un micropropulseur repose sur l'initiation thermique d'un matériau pyrotechnique de type propergol introduit dans la cavité des micropropulseurs. Une fois soumis à une polarisation de type courant, une résistance micro-usinée sur une membrane diélectrique très fine chauffe le propergol par effet Joule jusqu'à initié de l'auto-combustion. Les gaz générés vont traverser la micro-tuyère et fournir la poussée. Après avoir évalué les besoins en propulsion pour la stabilisation d'un drone miniature en vol, nous avons opté pour la micropropulsion à propergol solide qui présente de nombreux avantages pour l'application visée : c'est une technologie simple, nécessitant peu de puissance de fonctionnement (quelque 100mW) et qui est adaptable facilement au besoin de la mission. Les forces générées sont réglables entre quelques 100µN jusqu'au N en modifiant seulement - pour un type utilisé de propergol - la dimension du col de tuyère. Au cours de ce manuscrit de thèse, nous présenterons tout d'abord les spécifications de la DGA qui ont guidées nos conceptions, nous présenterons ensuite la technologie de fabrication et d'assemblage mis en œuvre au sein de la centrale technologique du LAAS. Et en fin, les résultats de caractérisation qui valident le fonctionnement et la gamme de poussée accessible par cette technologie seront donnés.

The CubeSat community has begun to develop and implement propulsion systems. This movement represents a new capability which may satisfy mission needs such as orbital and constellation maintenance, formation flight, de-orbit, and even interplanetary travel. With the freedom and capability granted by propulsion systems, CubeSat providers must accept new responsibilities in proportion to the potential hazards that propulsion systems may present. The Cal Poly CubeSat program publishes and maintains the CubeSat Design Specification (CDS). They wish to help the CubeSat community to safety and responsibly expand its capabilities to include propulsive designs. For this reason, the author embarked on the task of developing a draft of safety standards CubeSat propulsion systems. Wherever possible, the standards are based on existing documents. The author provides an overview of certain concepts in systems safety with respect to the classification of hazards, determination of required fault tolerances, and the use of inhibits to satisfy fault tolerance requirements. The author discusses hazards that could exist during ground operations and through launch with respect to hazardous materials and pressure systems. Most of the standards related to Range Safety are drawn from AFSPCMAN 91-710. Having reviewed a range of hypothetical propulsion system architectures with an engineer from Range Safety at Vandenberg Air Force Base, the author compiled a case study. The author discusses many aspects of orbital safety. The author discusses the risk of collision with the host vehicle and with third party satellites along with the trackability of CubeSats using propulsion systems. Some recommendations are given for working with the Joint Functional Component Command for Space (JFCC SPACE), thanks to the input of two engineers who work with the Joint Space Operations Center (JSpOC). Command Security is discussed as an important aspect of a mission which implements a propulsion system. The author also discusses End-of-Life procedures such as safing and de-orbit operations. The orbital safety standards are intended to promote “good citizenship.” The author steps through each proposed standard and offers justification. The author is confident that these standards will set the stage for a dialogue in the CubeSat community which will lead to the formulation of a reasonable and comprehensive set of standards. The author hopes that the discussions given throughout this document will help CubeSat developers to visualize the path to flight readiness so that they can get started on the right foot.