Colloid thrusters, and electrospray thrusters as a whole, have been around since the 1960s. When they were first developed, the high efficiency and fine thrust control was overshadowed by the high power requirement for such a low thrust that the system provides. This caused the technology to be put on hold for aerospace applications. Now, as small satellites are becoming more prevalent, there has been a resurgence in interest in electrospray thruster technology. The recent advancements in tech- nology allow electrospray thrusters to use significantly less power and occupy less volume than their predecessors. As electrospray technology continues to advance, these thrusters are meeting the demands of small satellite propulsion. As such, in an effort to keep the spacecraft propulsion curriculum current with today’s technology, a colloid thruster is designed, built, tested, and implemented as a laboratory activity at California Polytechnic State University, San Luis Obispo. Electrospray thrusters work by placing a voltage on an ionic liquid and extracting either beads of propellant or ions to generate thrust. By definition, colloid thrusters are a specific class of electrospray thrusters that use solvents, such as glycerol or formamide, to emit droplets or, in special cases, ions to generate thrust. To keep with the University’s “Learn by Doing” pedagogical philosophy, the thruster for this activity is designed to have a tactile and experiential impact on the students. The final design is a scaled up configuration of an existing electrospray design so that the students can easily see each component with the naked eye and can be correlated to a real world thruster that they might see in industry. As a laboratory experiment, the thruster needs to be able to utilize current equip- ment in the Space Environments and Testing Laboratory. One of the Student Vacuum Chambers (SVC) is utilized as well as two 1 kV power supplies and a 100V power supply. An indirect method of measuring performance metrics needs to be developed as there are no thrust balances sensitive enough in the lab designated for undergrad- uate use. As such, the students will be using the mass of the propellant, the time of operation, and knowledge of the propellant’s properties to estimate the performance of the thruster. To prove success of the thruster, a performance profile of the thruster is produced using an indirect method of measurement as well as visual observations of the thruster moving propellant byway of the electrospray theory. The tests show thrusts produced between 96-311 μN with an Isp ranging from 1270-1684 seconds. The visual evidence demonstrates propellant being collected as well as the operation of the thruster under the electrospray theory. The visual evidence also sheds light on which emission mode the thruster is operating at as well as a self-correcting failure mode that was occurring. The thruster is implemented as a lab for Cal Poly’s AERO 402 Spacecraft Propulsion Lab in Fall 2018, and it receives positive feedback from the students through an anonymous survey. While the colloid thruster demonstrates success in meeting performance and pedagog- ical goals, future work should be continued to improve the thruster. Further design and manufacturing work can be undertaken to improve the efficiency and decrease failure due to propellant impingement. Additionally, the procurement of power sup- plies capable of applying higher voltages can provide a greater range of operation which can enable a more dynamic student discovery of electrospray thrusters.

The use of Remotely Operated Vehicles (ROVs) in exploring and building infrastructure in the ocean is expanding. ROVs are performing tasks underwater that would be difficult or impossible to do with human divers. These vehicles are being used in increasingly complicated and demanding environments that require improvements in the methods for controlling these vehicles. Currently, research into semi-autonomous control is being conducted to aide ROV pilots in compensating for environmental disturbances and unknown dynamics. To effectively implement semi-autonomous control, precise thrust forces must be elicited from the thrusters. This work discusses a low-level thruster controller that can be used as part of a semi- autonomous guidance, navigation and control system for a ROV. A thruster dynamics model describing the thrust force of a propeller-type underwater thruster was derived and implemented for the thruster on the Saab Seaeye Falcon ROV. The thruster dynamics model described is a quadratic equation that uses the propeller velocity to determine thrust force. This model includes a mechanism for compensation against the external motion of the thruster, such as occurs when the ROV moves through the water. Several experiments were performed to fully characterize the quadratic thruster dynamics model and test its ability to accurately predict thrust force based on a known ambient water velocity and propeller angular velocity. The drag force was calculated and removed from the force measurements to get the thrust force used in the model. The model coefficients were determined and then the resulting model was tested against experimental data to determine the efficacy of the model in the lab environment and compare it to a widely used linear thruster dynamics model. The results showed the quadratic model improved upon the linear model, and the quadratic model was valid over a larger range of ambient water velocities. The quadratic model was then inverted to provide a thruster control algorithm that determines the propeller angular velocity necessary to produce a desired thrust force. This algorithm was used to design a low-level thruster controller. This controller was designed to be used on an existing vehicle where thrust force feedback is not available and difficult or expensive to add. This allows it to be used in a wider range of applications than controllers that rely on such feedback to operate. The controller was implemented using a PID control loop to drive the angular velocity of the propeller to the desired rate. An iso-parametric mapping, which transforms the linear PID output to the non-linear thruster input, was added to provide a faster response time for the controller over the entire range of the propeller velocity. The performance of this low-level thruster controller was demonstrated in the test environment. The low-level thruster controller followed a desired thrust force under a range of ambient water velocities. The thruster characterization and low-level thruster controller was designed to be used on an existing ROV. The motivation behind this work is to build a controller that may be implemented for use by a high-level vehicle controller. The low-level thruster controller presented here does not depend on sensors or equipment that is largely unavailable on vehicles without costly retrofits, and the experimental characterization does not require intimate knowledge of the inner workings of the thruster. This makes it easy to implement and generalize to a variety of thrusters. The results of this work show a low-level thruster controller than can be used in a control schema for existing ROVs. / Graduate / 0547 / matt@amosbuchanan.net

High Efficiency Multistage Plasma Thrusters (HEMPTs) are a relatively new form of electric propulsion that show promise for use on a variety of missions and have several advantages over their older EP competitors. One such advantage is their long predicted lifetime and minimal wall erosion due to a unique periodic permanent magnet system. A laboratory HEMPT was built and donated by JPL for testing at Cal Poly. Previous work was done to characterize the performance of this thruster and it was found to exhibit a large plume divergence, resulting in decreased thrust and specific impulse. This thesis explores the design and application of a magnetic shield to modify the thruster’s magnetic field to force more ion current towards the centerline. A previous Cal Poly thesis explored the same concept, and that work is continued and furthered here. The previous thesis tested a shield which increased centerline current but decreased performance. A new shield design which should avoid this performance decrease is studied here. Magnetic modelling of the thruster was performed using COMSOL. This model was verified using guassmeters to measure the field strength at many discrete points within and near the HEMPT, with a focus on the ionization channel and exit plane. A shield design which should significantly reduce the radial field strength at the exit plane without affecting the ionization channel field was modelled and implemented. The HEMPT was tested in a vacuum chamber with and without the shield to characterize any change to performance characteristics. Data were collected using a nude Faraday probe and retarding potential analyzer. The data show a significant increase in centerline current with the application of the shield, but due to RPA malfunction and thruster failure the actual change in performance could not be concluded. The unshielded HEMPT was characterized, however, and was found to produce 12.1 +/- 1.3 mN of thrust with a specific impulse of 1361 +/- 147s. The thruster operated with a total efficiency of 10.63 +/- 3.66%, an efficiency much lower than expected. A large contributor to this low efficiency is likely the use of argon in place of xenon. Its lower mass and higher ionization energy make it a less efficient propellant choice. Further, the thruster is prone to overheating, indicating that significant thermal losses are present in this design.

No description available.

This thesis discusses the recent developments to the electric propulsion plume code DRACO as well as a validation and sensitivity analysis of the code using data from an AFRL experiment using a Busek 200 W Hall Thruster. DRACO is a PIC code that models particles kinematically while using finite differences schemes to solve the electric potential and field. The DRACO code has been recently modified to improve simulation results, functionality and performance. A particle source has been added that uses the Hall Thruster device code HPHall as input for a source to model Hall Thrusters. The code is now also capable of using a non-uniform mesh that uses any combination of uniform, linear and exponential stretching schemes in any of the three directions. A stretched mesh can be used to refine simulation results in certain areas, such as the exit of a thruster, or improve performance by reducing the number of cells in a mesh. Finally, DRACO now has the capability of using a DSMC collision scheme as well as performing recombination collisions. A sensitivity analysis of the newly upgraded DRACO code was performed to test the new functionalities of the code as well as validate the code using experimental data gathered at AFRL using a Busek 200 W Hall Thruster. A simulation was created that attempts to numerically recreate the AFRL experiment and the validation is performed by comparing the plasma potential, polytropic temperature, ion number density of the thruster plume as well as Faraday and ExB probe results. The study compares the newly developed HPHall source with older source models and also compares the variations of the HPHall source. The field solver and collision model used are also compared to determine how to achieve the best results using the DRACO code. Finally, both uniform and non-uniform meshes are tested to determine if a non-uniform mesh can be properly implemented to improve simulation results and performance. The results from the validation and sensitivity study show that the DRACO code can be used to recreate a vacuum chamber simulation using a Hall Thruster. The best results occur when the newly developed HPHall source is used with a MCC collision scheme using a projected background neutral density and CEX collision tracking. A stretched mesh was tested and proved results that are as accurate as a uniform mesh, if not more accurate in locations of high mesh refinement. / Master of Science

Le propulseur PEGASES (Plasma Propulsion with Electronegative Gases) est un nouveau type de propulseur électrique pour la propulsion spatiale. Il utilise des ions négatifs et positifs créés par une décharge radiofréquence à couplage inductif pour générer la poussée. L’accélération électrostatique des ions est assurée par un ensemble de grilles polarisées. Un filtre magnétique est utilisé pour augmenter la quantité d'ions négatifs dans la cavité du propulseur. Le propulseur PEGASES est non seulement une source qui permet de créer un plasma d'ions négatifs à forte densité, et même un plasma d'ion-ion, mais il peut également être utilisé comme un propulseur ionique classique. Cela signifie qu'un plasma est créé dans un gaz électropositif (e.g. Xe) et que les ions positifs sont extraits et accélérés. Dans ce cas, il est nécessaire de neutraliser le plasma derrière la zone d'accélération, comme dans d'autres propulseurs ioniques. Les performances du propulseur PEGASES ont été étudiées principalement dans du xénon afin de comparer les résultats obtenus avec les propulseurs ioniques de type RIT. Le propulseur a été étudié à l'aide d'une série de sondes telles qu’une sonde de Langmuir, une sonde plane, une sonde capacitive et un RPA (pour Analyseur à Champ Retardateur). De plus, une sonde en champs croisés ExB a été développée pour mesurer la vitesse des ions quittant le propulseur ainsi que la fraction des différentes espèces ioniques présentes dans le plasma. / The PEGASES thruster (Plasma Propulsion with Electronegative Gases) is a novel type of electric thruster for space propulsion. It uses negative and positive ions produced by an inductively coupled radio frequency discharge to create the thrust by electrostatically accelerating the ions through a set of grids. A magnetic filter is used to increase the amount of negative ions in the cavity of the thruster. The PEGASES thruster is not only a source to create a strongly negative ion plasma or even an ion-ion plasma but it can also be used as a classical ion thruster. This means that a plasma is created and only the positive ions are extracted and accelerated making it necessary to neutralize the plasma behind the acceleration stage like in other ion thrusters. The performances of the PEGASES thruster have been investigated mainly in xenon in order to compare the obtained results with RIT-type ion thrusters. The thruster has been investigated with the help of a variety of probes such as a Langmuir probe, a planar probe, a capacitive probe and a RPA (Retarding Potential Analyzer). In addition, an ExB probe has been developed to measure the velocity of the ions leaving the thruster and to differentiate between the ion species present in the plasma.

abstract: This thesis gives a detailed design process for a pulsed type thruster. The thrust stand designed in this paper is for a Pulsed Plasma Thruster built by Sun Devil Satellite Laboratory, a student organization at Arizona State University. The thrust stand uses a torsional beam rotating to record displacement. This information, along with impulse-momentum theorem is applied to find the impulse bit of the thruster, which varies largely from other designs which focus on using the natural dynamics their fixtures. The target impulse to record on this fixture was estimated to be 275 μN-s of impulse. Through calibration and experimentation, the fixture is capable of recording an impulse of 332 μN-s ± 14.81 μN-s, close to the target impulse. The error due to noise was characterized and evaluated to be under 5% which is deemed to be acceptable. / Dissertation/Thesis / Masters Thesis Aerospace Engineering 2017

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:

Vacuum arc thrusters offer a relatively simple and cheap form of satellite propulsion, especially suitable for nanosatellites such as CubeSats or even smaller. This thesis focuses on vacuum arc thruster design considering the thruster’s manufacturing, assembly and integration into the spacecraft, and proposes a new anode geometry easing thruster production. Vacuum arc thruster research is traditionally experimental in nature due to a lack of accurate models. This work follows this approach, and studies experimentally the effect of several geometric design parameters on thruster performance. The outcome confrms findings from several papers, and suggests specifc improvements towards existing models for predicting the effect of the thruster’s geometry on its thrust. The chosen experimental approach raised the need for a micro-thrust measurement stand. Two distinct measurement stands have been designed, realized and used to test various thruster prototypes. One test stand is more accurate. However, the other setup allows for considerably faster testing.

A new particle-in-cell algorithm was developed for plasma simulations involving complex boundary conditions. The new algorithm is based on the three-dimensional immersed finite element method which is developed in this thesis, and a modified legacy particle-in-cell code. The model also applies a new meshing technique that separates the field solution mesh from the particle pushing mesh in order to increase the computational eciency of the model. The new simulation model is used in two applications of great importance to the development of ion propulsion technology: the ion optics performance and the interaction between spacecraft and the ion thruster. The first application is ion optics simulations. Simulations are performed to investigate ion optics plasma flow for a whole subscale NEXT ion optics. The operating conditions modeled cover the entire cross-over to perveance limit range. The results of the ion optics simulations demonstrated good agreement with the available experimental data. The second application is ion thruster plume simulations. Simulations are performed to investigate ion thruster plume - spacecraft interactions for the Dawn spacecraft. Plume induced contaminations on the solar array are studied for a variety of ion thruster configurations including multiple thruster firings. / Ph. D.