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.

Thesis. 1978. M.S.--Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERONAUTICS. / Includes bibliographical references. / M.S.

Thesis. 1978. M.S.--Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics. / This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERONAUTICS. / Includes bibliographical references. / M.S.

The understanding and prediction of transient phenomena inside Liquid Rocket Engines (LREs) have been very difficult because of the many challenges posed by the conditions inside the combustion chamber. This is especially true for injectors involving liquid oxygen LOX and gaseous hydrogen GH₂. A wide range of length scales needs to be captured from high-pressure flame thicknesses of a few microns to the length of the chamber of the order of a meter. A wide range of time scales needs to be captured, again from the very small timescales involved in hydrogen chemistry to low-frequency longitudinal acoustics in the chamber. A wide range of densities needs to be captured, from the cryogenic liquid oxygen to the very hot and light combustion products. A wide range of flow speeds needs to be captured, from the incompressible liquid oxygen jet to the supersonic nozzle. Whether one desires to study these issues numerically or experimentally, they combine to make simulations and measurements very difficult whereas reliable and accurate data are required to understand the complex physics at stake. This thesis focuses on the numerical simulations of flows relevant to LRE applications using Large Eddy Simulations (LES). It identifies the required features to tackle such complex flows, implements and develops state-of-the-art solutions and apply them to a variety of increasingly difficult problems. More precisely, a multi-species real gas framework is developed inside a conservative, compressible solver that uses a state-of-the-art hybrid scheme to capture at the same time the large density gradients and the turbulent structures that can be found in a high-pressure liquid rocket engine. Particular care is applied to the implementation of the real gas framework with detailed derivations of thermodynamic properties, a modular implementation of select equations of state in the solver. and a new efficient iterative method. Several verification cases are performed to evaluate this implementation and the conservative properties of the solver. It is then validated against laboratory-scaled flows relevant to rocket engines, from a gas-gas reacting injector to a liquid-gas injector under non-reacting and reacting conditions. All the injectors considered contain a single shear coaxial element and the reacting cases only deal with H₂-O₂ systems. A gaseous oyxgen-gaseous hydrogen (GOX-GH₂) shear coaxial injector, typical of a staged combustion engine, is first investigated. Available experimental data is limited to the wall heat flux but extensive comparisons are conducted between three-dimensional and axisymmetric solutions generated by this solver as well as by other state-of-the-art solvers through a NASA validation campaign. It is found that the unsteady and three-dimensional character of LES is critical in capturing physical flow features, even on a relatively coarse grid and using a 7-step mechanism instead of a 21-step mechanism. The predictions of the wall heat flux, the only available data, are not very good and highlight the importance of grid resolution and near-wall models for LES. To perform more quantitative comparisons, a new experimental setup is investigated under both non-reacting and reacting conditions. The main difference with the previous setup, and in fact with most of the other laboratory rigs from the literature, is the presence of a strong co-flow to mimic the surrounding flow of other injecting elements. For the non-reacting case, agreement with the experimental high-speed visualization is very good, both qualitatively and quantitatively but for the reacting case, only poor agreement is obtained, with the numerical flame significantly shorter than the observed one. In both cases, the role of the co-flow and inlet conditions are investigated and highlighted. A validated LES solver should be able to go beyond some experimental constraints and help define the next direction of investigation. For the non-reacting case, a new scaling law is suggested after a review of the existing literature and a new numerical experiment agrees with the prediction of this scaling law. A slightly modified version of this non-reacting setup is also used to investigate and validate the Linear-Eddy Model (LEM), an advanced sub-grid closure model, in real gas flows for the first time. Finally, the structure of the trans-critical flame observed in the reacting case hints at the need for such more advanced turbulent combustion model for this class of flow.

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.

Future ambitious solar system exploration missions are likely to require ever larger propulsion capabilities and involve innovative interplanetary trajectories in order to accommodate the increasingly complex mission scenarios. Two recent advances in trajectory design can be exploited to meet those new requirements: the use of low-thrust propulsion which enables larger cumulative momentum exchange relative to chemical propulsion; and the consideration of low-energy transfers relying on full multi-body dynamics. Yet the resulting optimal control problems are hypersensitive, time-consuming and extremely difficult to tackle with current optimization tools. Therefore, the goal of the thesis is to develop a methodology that facilitates and simplifies the solution finding process of low-thrust optimization problems in multi-body environments. Emphasis is placed on robust techniques to produce good solutions for a wide range of cases despite the strong nonlinearities of the problems. The complete trajectory is broken down into different component phases, which facilitates the modeling of the effects of multiple bodies and makes the process less sensitive to the initial guess. A unified optimization framework is created to solve the resulting multi-phase optimal control problems. Interfaces to state-of-the-art solvers SNOPT and IPOPT are included. In addition, a new, robust Hybrid Differential Dynamic Programming (HDDP) algorithm is developed. HDDP is based on differential dynamic programming, a proven robust second-order technique that relies on Bellman's Principle of Optimality and successive minimization of quadratic approximations. HDDP also incorporates nonlinear mathematical programming techniques to increase efficiency, and decouples the optimization from the dynamics using first- and second-order state transition matrices. Crucial to this optimization procedure is the generation of the sensitivities with respect to the variables of the system. In the context of trajectory optimization, these derivatives are often tedious and cumbersome to estimate analytically, especially when complex multi-body dynamics are considered. To produce a solution with minimal effort, an new approach is derived that computes automatically first- and high-order derivatives via multicomplex numbers. Another important aspect of the methodology is the representation of low-thrust trajectories by different dynamical models with varying degrees of fidelity. Emphasis is given on analytical expressions to speed up the optimization process. In particular, one novelty of the framework is the derivation and implementation of analytic expressions for motion subjected to Newtonian gravitation plus an additional constant inertial force. Example applications include low-thrust asteroid tour design, multiple flyby trajectories, and planetary inter-moon transfers. In the latter case, we generate good initial guesses using dynamical systems theory to exploit the chaotic nature of these multi-body systems. The developed optimization framework is then used to generate low-energy, inter-moon trajectories with multiple resonant gravity assists.

This work focuses on improving the thrust-to-power ratio of Hall effect thrusters using in-channel electrodes to reduce ion-wall neutralization and focus the ion beam. A higher thrust-to-power ratio would give Hall thrusters increased thrust with the limited power available on spacecraft. A T-220HT Hall thruster is modified in this work to include a pair of ring electrodes within inside the discharge channel. The electrodes are biased above anode potential to repel ions from the walls and toward the channel centerline. Theoretical analysis of ion loss factors indicate that ion-wall neutralizations remove almost 13% of the total ions produced. Reduced wall losses could significantly improve the thruster performance without increased discharge power or propellant consumption. The thruster performance, plume ion characteristics, and internal plasma contours are experimentally measured. The plume and internal plasma measurements are important to determine the cause of the performance changes. The thruster is tested in three conditions: no electrode bias, low bias (10 V), and high bias (30 V). The performance measurements show the electrodes do indeed improve the thrust and thrust-to-power ratio, the latter only at the low bias level. Adding bias increases the ion density and decreases the plume angle compared to the no bias case. The plume measurements indicate that the performance improvements at low bias are due to increased ion number density as opposed to increased ion energy. The increased ion density is attributed to reduced wall losses, not increased ionization. The in-channel measurements support this due to little change in the acceleration potential or the electron temperature. At the high bias level, a drop in thrust-to-power ratio is seen, even though a larger increase in thrust is observed. This is due to increased power draw by the electrodes. Plume measurements reveal the increased thrust is due to ion acceleration. The internal measurements show increased acceleration potential and electron energy which can lead to increased ionization. At the high bias condition, the electrodes become the dominant positive terminal in the thruster circuit. This causes the increased ion acceleration and the creation of domed potential contours that conform to the near-wall cusp-magnetic fields. The domed contours produce focused electric fields, which cause the decreased wall losses and plume angle.

For the past four decades, Mars missions have relied on Viking heritage technology for supersonic descent. Extending the use of propulsion, which is required for Mars subsonic deceleration, into the supersonic regime allows the ability to land larger payload masses. Wind tunnel and computational experiments on subscale supersonic retropropulsion models have shown a complex aerodynamic flow field characterized by the interaction of underexpanded jet plumes exhausting from nozzles on the vehicle with the supersonic freestream. Understanding the impact of vehicle and nozzle configuration on this interaction is critical for analyzing the performance of a supersonic retropropulsion system, as deceleration will have components provided by both the aerodynamic drag of the vehicle and thrust from the nozzles. This investigation focuses on the validity of steady state computational approaches to analyze supersonic retropropulsion flowfield structures and their effect on vehicle aerodynamics. Wind tunnel data for a single nozzle and a multiple nozzle configuration are used to validate a steady state, turbulent computational fluid dynamics approach to modeling supersonic retropropulsion. An analytic approximation to determine plume and bow shock structure in the flow field is also developed, enabling rapid assessment of flowfield structure for use in improved grid generation and as a configuration screening tool. Results for both the computational fluid dynamics and analytic approaches show good agreement with the experimental datasets. Potential limitations of the two methods are identified based on the comparisons with available data. Six additional geometries are defined to investigate the extensibility of the analytical model and determine the variation of supersonic retropropulsion performance with configuration. These validation geometries are split into two categories: three geometries with nozzles located on the vehicle forebody at varying nozzle cant angles, and three geometries with nozzles located on the vehicle aftbody at varying nozzle cant angles and number of nozzles. The forebody nozzle configurations show that nozzle cant angle is a significant driver in performance of a vehicle employing supersonic retropropulsion. Aerodynamic drag preservation for a given thrust level increases with increasing cant angle. However, increasing the cant angle reduces the contribution of thrust to deceleration. The tradeoff between these two contributions to the deceleration force is examined, noting that performance improvements are possible with modest nozzle cant angles. Static pitch stability characteristics are investigated for the lowest and highest cant angle configurations. The aftbody nozzle configuration results show that removing the plume flow from the region forward of the vehicle results in less interaction with the bow shock structure. This impacts aerodynamic performance, as the surface pressure remains relatively undisturbed for all thrust values examined. Static pitch stability characteristics for each of the aftbody nozzle configurations are investigated; noting that supersonic retropropulsion for these configurations exhibits a transition point from static stability to instability as a function of this center of mass location along the axis.

An extension of the current acetone Planar Laser-Induced Fluorescence (PLIF) was formulated for mixing studies of fluids at subcritical and supercritical conditions. The new technique, called Planar Laser-Induced Fluorescence and Phosphorescence (PLIFP), employs the difference in the mass diffusivity of the denser (liquid) to the less dense (vapor/supercritical fluid) to delineate the interface where a phase change occurs. The vapor/supercritical acetone fluorescence signal is utilized to measure of the acetone vapor density, the mixture fractions and liquid acetone phosphorescence signal to determine the location of the phase interface. The application of the technique requires the photophysical properties of vapor and liquid acetone to be known. Therefore, a series of controlled experiments were done to determine their photophysics at elevated temperatures and pressures up to T/TC = 1.2 and p/pC =1.25. The demonstration of the techniques shows it was able to provide quantitative measurements of acetone number density and the overall mixture fraction within the test chamber. Also, the size and mass of droplets that have broken off from the main jet were determined as well, though the ability is limited to small droplets (d~100 μm). In addition, the technique was able to delineate the low diffusivity (subcritical)/high diffusivity (supercritical) interface very well.