The rectangular pulsed plasma thruster (PPT) is an electromagnetic thruster that ablates Teflon propellant to produce thrust in a discharge that lasts 5-20 microseconds. In order to integrate PPTs onto spacecraft, it is necessary to investigate possible thruster plume-spacecraft interactions. The PPT plume consists of neutral and charged particles from the ablation of the Teflon fuel bar as well as electrode materials. In this thesis a novel application of quadruple Langmuir probes is implemented in the PPT plume to obtain electron temperature, electron density, and ion speed ratio measurements (ion speed divided by most probable thermal speed). The pulsed plasma thruster used is a NASA Glenn laboratory model based on the LES 8/9 series of PPTs, and is similar in design to the Earth Observing-1 satellite PPT. At the 20 J discharge energy level, the thruster ablates 26.6 mg of Teflon, creating an impulse bit of 256 mN-s with a specific impulse of 986 s. The quadruple probes were operated in the so-called current mode, eliminating the need to make voltage measurements. The current collection to the parallel to the flow electrodes is based on Laframboise's theory for probe to Debye length ratios between 5 and 100, and on the thin-sheath theory for ratios above 100. The ion current to the perpendicular probe is based on a model by Kanal and is a function of the ion speed ratio, the applied non-dimensional potential and the collection area. A formal error analysis is performed using the complete set of nonlinear current collection equations. The quadruple Langmuir probes were mounted on a computer controlled motion system that allowed movement in the radial direction, and the thruster was mounted on a motion system that allowed angular variation. Measurements were taken at 10, 15 and 20 cm form the Teflon fuel bar face, at angles up to 40 degrees off of the centerline axis at discharge energy levels of 5, 20, and 40 J. All data points are based on an average of four PPT pulses. Data analysis shows the temporal and spatial variation in the plume. Electron temperatures show two peaks during the length of the pulse, a trend most evident during the 20 J and 40 J discharge energies at 10 cm from the surface of the Teflon fuel bar. The electron temperatures after the initial high temperature peak are below 2 eV. Electron densities are highest near the thruster exit plane. At 10 cm from the Teflon surface, maximum electron densities are 1.04e20 ± 2.8e19 m-3, 9.8e20 ± 2.3e20 m-3, and 1.38e21 ± 4.05e20 m-3 for the 5 J, 20 J and 40 J discharge energy, respectively. The electrons densities decrease to 2.8x1019 ± 8.9e18 m-3, 1.2e20 ± 4.2e19 m-3, and 4.5e20 ± 1.2e20 m-3 at 20 cm for the 5 J, 20 J, and 40 J cases, respectively. Electron temperature and density decrease with increasing angle away from the centerline, and with increasing downstream distance. The plume is more symmetric in the parallel plane than in the perpendicular plane. Ion speed ratios are lowest near the thruster exit, increase with increasing downstream distance, but do not show any consistent angular variation. Peak speed ratios at a radial distance of 10 cm are 5.9±3.6, 5.3±0.39, and 4.8±0.41 for the 5 J, 20 J and 40 J discharge energies, respectively. The ratios increase to 6.05±5.9, 7.5±1.6, and 6.09±0.72 at a radial distance of 20 cm. Estimates of ion velocities show peak values between 36 km/s to 40 km/s, 26 km/s to 30 km/s, and 26 km/s to 36 km/s for the % J, 20 J, and 40 J discharge energies, respectively.

Colloid thrusters are under consideration for NASA missions such as the Laser Interferometer Space Antenna (LISA), which requires the continuous cancellation of external disturbances (approximately 25 microNewtons over a 3-10 year mission). Emissive probes are one diagnostic for the measurement of plasma potential, which can provide valuable information on the level of space-charge neutralization in a thruster plume. Understanding how to achieve effective space-charge neutralization of the positive-droplet thruster plume is important for efficient operation and to minimize the risk of contamination. In this Thesis we describe a laboratory electrospray (colloid) source and accompanying power processing electronics developed for testing of diagnostics in colloid thruster plumes. We present results of an initial series of emissive probe measurements using floating probe and swept bias probe techniques. These measurements were carried out using a single needle emitter operating on a mixture of EMI-IM (an ionic liquid) and tributyl phosphate. For a spray operating at a discharge voltage and current of 2.0kV and 200nA respectively, a potential of 5.0V was measured using the floating probe technique with the probe located at a distance of 2.7cm from the electrospray source. The interpretation of this floating potential as the plasma potential is discussed. In a separate set of tests, we used the swept bias emissive probe technique at the same distance and measured a plasma potential of 2.0V at a discharge voltage of 2.0kV. The discharge current in this latter test was somewhat unstable and varied from approximately 250 nA to over 1000nA. Numerical integration of the Poisson equation was performed to better understand space charge limitations of a probe emitting into a low density plasma. These results are presented and some implications for the measurements discussed. While the electrospray droplet number density was not measured, calculations to estimate this number density are also presented. Based on these estimates and our numerical calculations, the“knee" in the current voltage characteristic measured using the swept probe technique is estimated to be within 1.3 V of the actual plasma potential.

Submitted by Maria Beatriz Vieira (mbeatriz.vieira@gmail.com) on 2017-05-29T15:07:15Z No. of bitstreams: 2 license_rdf: 0 bytes, checksum: d41d8cd98f00b204e9800998ecf8427e (MD5) dissertacao_bruno_kabke_bainy.pdf: 1791571 bytes, checksum: fbfdf2a30e93aa9c1820193c665cd912 (MD5) / Approved for entry into archive by Aline Batista (alinehb.ufpel@gmail.com) on 2017-05-29T21:22:58Z (GMT) No. of bitstreams: 2 license_rdf: 0 bytes, checksum: d41d8cd98f00b204e9800998ecf8427e (MD5) dissertacao_bruno_kabke_bainy.pdf: 1791571 bytes, checksum: fbfdf2a30e93aa9c1820193c665cd912 (MD5) / Made available in DSpace on 2017-05-29T21:22:58Z (GMT). No. of bitstreams: 2 license_rdf: 0 bytes, checksum: d41d8cd98f00b204e9800998ecf8427e (MD5) dissertacao_bruno_kabke_bainy.pdf: 1791571 bytes, checksum: fbfdf2a30e93aa9c1820193c665cd912 (MD5) Previous issue date: 2015-02-24 / Coordenação de Aperfeiçoamento de Pessoal de Nível Superior - CAPES / Esta dissertação de mestrado propõe a elaboração inicial de um modelo para a dispersão de efluentes de foguetes e veículos espaciais. Neste estudo foi desenvolvida uma solução para a equação de advecção- difusão bidimensional transiente através da técnica GILTT, além de terem sido compilada da literatura algumas formulações para parâmetros micrometeorológicos e outras variáveis que representam fenômenos relevantes nas atividades de lançamento de foguetes. O modelo de dispersão foi testado com os experimentos de Hanford e Copenhagen com ótimos resultados. Além disso, foi rodado um caso particular para a região do Centro de Lançamentos de Alcântara para exemplificar e apresentar maiores detalhes do modelo. / This master thesis proposes a first attempt to elaborate a model for rocket exhaust dispersion. In this study, a solution to the time-dependant two-dimensional advectiondiffusion equation was obtained through the GILTT, as well as it assembles of some literature formulations for micrometeorological parameters and other variables which represent important phenomena in space vehicles launching. The dispersion model was tested against two experimental data, Hanford and Copenhagen, with great results, and an additional simulation was run using data from the Alcantara Launch Centre, aiming to exemplify and present aditional details of the model.

During hypersonic atmospheric entry, spacecraft are exposed to enormous aerodynamic heat. To prevent the payload from overheating, charring ablative materials are favored to be applied as the heat shield at the exposing surface of the vehicle. Accurate modeling not only prevents mission failures, but also helps reduce cost. Existing models were mostly limited to one-dimensional and discrepancies were shown against measured experiments and flight-data. To help improve the models and analyze the charring ablation problems, a multidimensional material response module is developed, based on a finite volume method framework. The developed computer program is verified through a series of test-cases, and through code-to-code comparisons with a validated code. Several novel models are proposed, including a three-dimensional pyrolysis gas transport model and an orthotropic material model. The effects of these models are numerically studied and demonstrated to be significant.

The goal of this thesis is to use the Cal Poly Spacecraft Attitude Dynamics Simulator to provide proof of concept of two adaptive control theories developed by former Cal Poly students: Nonlinear Direct Model Reference Adaptive Control and Adaptive Output Feedback Control. The Spacecraft Attitude Dynamics Simulator is a student-built air bearing spacecraft simulator controlled by four reaction wheels in a pyramidal arrangement. Tests were performed to determine the effectiveness of the two adaptive control theories under nominal operating conditions, a “plug-and-play” spacecraft scenario, and under simulated actuator damage. Proof of concept of the adaptive control theories applied to attitude control of a spacecraft is provided. The adaptive control theories are shown to attain similar or improved performance over a Full State Feedback controller. However, the measurement capabilities of the simulator need to be improved before strong comparisons between the adaptive controllers and Full State Feedback can be achieved.

Most combustion machines feature internal flows with very high energy densities. If a small fraction of the total energy contained in the flow is diverted into oscillations, large mechanical or thermal loads on the structure can be the result, which are potentially devastating if not predicted correctly. This is particularly the case for lightweight high performing devices like rockets. The problem is commonly known as "Combustion Instability". Several mechanisms have been identified in the past that link the flow field to the acoustics inside a combustion chamber and thereby drive or dampen oscillations, one of them being vortex shedding. The interaction between the highly sheared flow behind an obstacle and longitudinal acoustic oscillations inside a solid rocket booster is investigated both analytically and experimentally.The analytical approach is developed based on modeling of the second order acoustic energy. The energy model is applied to the specific flow conditions just downstream of a single baffle protruding into the flow. The mean flow profile is assumed to be of the form of a hyperbolic tangent, the unsteady acoustic velocities are assumed to be sinusoidally oscillating. Solutions for the unsteady rotational velocities and the unsteady vorticity are derived. The resulting flow field is utilized in stability calculations for a simplified two-dimensional axial-symmetric geometry. This yields to linear growth rates of the (longitudinal) oscillation modes. The growth rates are functions of the chamber geometry, the mean flow properties and the properties of the shear layer created by the flow restriction.A cold flow experiment is designed, tested and performed in order to validate the analytical findings. Flow is injected radially into a tube with acoustic closed-closed end conditions. A single baffle is installed in the tube, the axial position of the baffle is varied as well as its inner diameter. Frequency spectra of pressure oscillations are recorded. The experimental data is then compared qualitatively to the analytical growth rates. Those longitudinal Normal Modes, which feature the highest theoretical growth rates, are expected to be most prominent in the experimental data. This behavior is clearly observable.

The space vehicle system concept (i.e. resupply vehicle) described is based on the new direction that President George W. Bush announced on January 14, 2004 for NASA’s Human Exploration, which has the space shuttle retiring in 2011 following the completion of the International Space Station (ISS). This leads to a problem for the ISS community regarding the capability of meeting a sixty metric-ton cargo shortfall in resupply and the ability of returning large payloads, experiment racks and any other items too large to fit into a crew only type spacecraft like the Orion or Soyuz. NASA and the ISS partners have realized these future problems and started developing various systems for resupply to ISS, but none offer the capability for large up or down mass close to that of the shuttle. Without this capability, the primary purpose behind the ISS science mission is defeated and the ability to keep the station functioning properly is at risk with limited payload delivery (i.e. replacement hardware size and mass). There is a solution to this problem and a majority of the solution has already been designed, built, and flight tested. Another portion has been studied heavily by a team at NASA for use in a slightly different mission. Following the retirement of the space shuttle fleet and the loss of heavy up and down mass capability, the only solution to the problem is to design a new spacecraft. However, the budget and new direction for NASA will not allow for a costly new payload carrying spacecraft. The solution is to use existing commercial off the shelf (COTS) hardware to minimize the costs of developing a totally new system. This paper will discuss the technical feasibility of this conceptual configuration.

During development of a complex system, feasibility initially overshadows other concerns, in some cases leading to a design which may not be viable long-term. In particular for the case of Reusable Launch Vehicles, Operations&Maintenance comprises the majority of the vehicle's LCC, whose stochastic nature precludes direct analysis. Through the use of simulation, probabilistic methods can however provide estimates on the economic behavior of such a system as it evolves over time. Here the problem of operations optimization is examined through the use of discrete event simulation. The resulting tool built from the lessons learned in the literature review simulates a RLV or fleet of vehicles undergoing maintenance and the maintenance sites it/they visit as the campaign evolves over a period of time. The goal of this work is to develop a method for uncovering an optimal operations scheme by investigating the effect of maintenance technician skillset distributions on important metrics such as the achievable annual flight rate and maintenance man hours spent on each vehicle per flight. Using these metrics, the availability of technicians for each subsystem is optimized to levels which produce the greatest revenue from flights and minimum expenditure from maintenance.

A study has been performed to determine the advantages and disadvantages of variable thrust and variable specific impulse (Isp) trajectories for solar system exploration. There have been several numerical research efforts for variable thrust, variable Isp, power-limited trajectory optimization problems. All of these results conclude that variable thrust, variable Isp (variable specific impulse, or VSI) engines are superior to constant thrust, constant Isp (constant specific impulse, or CSI) engines. However, most of these research efforts assume a mission from Earth to Mars, and some of them further assume that these planets are circular and coplanar. Hence they still lack the generality. This research has been conducted to answer the following questions: - Is a VSI engine always better than a CSI engine or a high thrust engine for any mission to any planet with any time of flight considering lower propellant mass as the sole criterion? - If a planetary swing-by is used for a VSI trajectory, is the fuel savings of a VSI swing-by trajectory better than that of a CSI swing-by or high thrust swing-by trajectory? To support this research, an unique, new computer-based interplanetary trajectory calculation program has been created. This program utilizes a calculus of variations algorithm to perform overall optimization of thrust, Isp, and thrust vector direction along a trajectory that minimizes fuel consumption for interplanetary travel. It is assumed that the propulsion system is power-limited, and thus the compromise between thrust and Isp is a variable to be optimized along the flight path. This program is capable of optimizing not only variable thrust trajectories but also constant thrust trajectories in 3-D space using a planetary ephemeris database. It is also capable of conducting planetary swing-bys. Using this program, various Earth-originating trajectories have been investigated and the optimized results have been compared to traditional CSI and high thrust trajectory solutions. Results show that VSI rocket engines reduce fuel requirements for any mission compared to CSI rocket engines. Fuel can be saved by applying swing-by maneuvers for VSI engines, but the effects of swing-bys due to VSI engines are smaller than that of CSI or high thrust engines.

Reservoir-type dispenser hollow cathodes are currently being developed for use on NASAs Prometheus 1 mission. In these cathodes, the reaction between a barium source material and tungsten powder contained in a cavity surrounding a porous tungsten emitter produces barium vapor which is crucial to operation of the cathode. The primary objective of this research was to investigate the reactions between tungsten and a commercial barium source material in a simulated reservoir hollow cath-ode environment. Mixtures of tungsten and a barium calcium aluminate material were sealed inside molybdenum capsules with porous tungsten closures and heated to 1000?1200?and 1300?or 100, 200, and 400 hours. Based on the reaction products, which were identified to be BaAl2O4 and Ba2CaWO6, a reaction was proposed for the barium calcium aluminate material with tungsten. The bottom pellets in the capsules were found to have reacted to a much further extent than the top pellets in all of the samples, possibly due to a temperature gradient or excessive moisture in the base of the capsules. Quantita-tive and semi-quantitative x-ray analysis results did not show a clear trend as to how the concentrations of BaAl2O4 and Ba2CaWO6 vary with time. Most of the barium source materials are hygroscopic, and hydration of the materi-als would substantially reduce the performance of the cathode. Therefore, the environ-mental stability of several barium compounds, 3BaO??2O3 (B3A), 6BaO????2O3 (612), 4BaO????O3 (411), Ba2.9Ca1.1Al2O7 (B4ASSL), and Ba3Sc4O9, were investi-gated in order to evaluate their suitability for use as barium source materials. A micro-balance was used to measure weight gain of the materials as they were exposed to dew points of ??C and 11?t room temperature. The results showed that B3A hydrated more extensively than any of the other materials tested in the low- and intermediate-humidity environments, while the 612, 411, and B4ASSL materials were all reasonably stable in the low-humidity environment. The Ba3Sc4O9 was extremely stable compared to the barium aluminates in the intermediate-humidity conditions.