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.

This Master's Thesis handles prestudy work and early hardware development that resulted in architectural definitions and prototype hardware of electronic ground support equipment. This equipment is destined to emulate the electric power consumption of the PPS5000 Hall Effect Thruster (HET), for use in satellite end-to-end tests of the all-electric Geostationary Satellite Electra, developed at OHB Sweden AB. The Thruster Emulator (TEM) was defined through a resulting compilation of intricate interdependent components that interface the satellite power system and the thruster, which yielded an architecture development to support some basic predefined emulator requirements. This architecture was then analyzed to form a base-line conceptual function of the emulator system, which incorporates the entire HET functionality. Six primary HET impedances were defined, of which the three most complex impedances were investigated fully. For the primary thruster discharge, research is shown of the complexity of implementing advanced electronic load hardware directly to the satellite's 5kW power system with respect to the transient primary plasma discharge during thruster start up, and with limitations on the electronic load reducing emulator-thruster similarities. Additionally, a fully functional plasma ignition emulator prototype circuit board was built to be used in the final hardware of the TEM to emulate the external HET cathode start-up functionality. Finally, a feasibility study for designing a possible solution for the large PPS5000 electromagnet impedance was performed, resulting in the manufacture of two prototype inductors with unsatisfying performance results according to the design requirements.

<p>The Earth's bow shock is a boundary where the solar wind becomes decelerated from supersonic to subsonic speed before being deflected around the Earth. This thesis presents measurements by the Cluster spacecraft upstream and at the Earth's quasi-parallel bow shock where the angle between the upstream magnetic field and the bow shock normal is less than 45 degrees. An intrinsic feature of quasi-parallel shocks is the ability of ions, that are reflected off the shock in a specular manner, to propagate far upstream and to interact with the incident solar wind. This leads to the generation of a variety of plasma waves, e.g., Ultra-Low Frequency (ULF) waves, which in their turn interact with the different ion populations. Some of the ULF waves are thought to steepen into so-called Short Large-Amplitude Magnetic Structures (SLAMS). </p><p>This thesis studies the impact of SLAMS on the incident solar wind. SLAMS are thought to play an important role in terms of 1) returning shock-reflected ions back to the shock where they can eventually contribute to downstream thermalisation and 2) local pre-dissipation of the solar wind. </p><p>The first electric field measurements of SLAMS showed a strong electric field rotation over SLAMS in association with the rotation of the magnetic field. This often leads to a local change from quasi-parallel to quasi-perpendicular conditions. In addition, short-scale electric field features were observed, e.g., spiky electric field structures associated with the leading edge of SLAMS and solitary electric field structures on Debye length scales, which are suggested to represent ion phase space holes. </p><p>Using the abilitiy of the four Cluster satellites to obtain propagation vectors of SLAMS and the high-resolution electric field measurements, the electric potential over SLAMS was studied. These structures are associated with a significant potential on the order of a few hundred to thousand Volt. Comparing these findings with data from the ion spectrometer, it was found that the bulk flow is locally significantly decelerated and moderately deflected and heated. In addition, SLAMS reflect incident ions on both the leading and trailing edge. The flux of so-called gyrating ions show a clear maximum in association with SLAMS. This indicates that SLAMS indeed play an important role for pre-dissipation of the solar wind upstream of the shock.</p>

The Earth's bow shock is a boundary where the solar wind becomes decelerated from supersonic to subsonic speed before being deflected around the Earth. This thesis presents measurements by the Cluster spacecraft upstream and at the Earth's quasi-parallel bow shock where the angle between the upstream magnetic field and the bow shock normal is less than 45 degrees. An intrinsic feature of quasi-parallel shocks is the ability of ions, that are reflected off the shock in a specular manner, to propagate far upstream and to interact with the incident solar wind. This leads to the generation of a variety of plasma waves, e.g., Ultra-Low Frequency (ULF) waves, which in their turn interact with the different ion populations. Some of the ULF waves are thought to steepen into so-called Short Large-Amplitude Magnetic Structures (SLAMS). This thesis studies the impact of SLAMS on the incident solar wind. SLAMS are thought to play an important role in terms of 1) returning shock-reflected ions back to the shock where they can eventually contribute to downstream thermalisation and 2) local pre-dissipation of the solar wind. The first electric field measurements of SLAMS showed a strong electric field rotation over SLAMS in association with the rotation of the magnetic field. This often leads to a local change from quasi-parallel to quasi-perpendicular conditions. In addition, short-scale electric field features were observed, e.g., spiky electric field structures associated with the leading edge of SLAMS and solitary electric field structures on Debye length scales, which are suggested to represent ion phase space holes. Using the abilitiy of the four Cluster satellites to obtain propagation vectors of SLAMS and the high-resolution electric field measurements, the electric potential over SLAMS was studied. These structures are associated with a significant potential on the order of a few hundred to thousand Volt. Comparing these findings with data from the ion spectrometer, it was found that the bulk flow is locally significantly decelerated and moderately deflected and heated. In addition, SLAMS reflect incident ions on both the leading and trailing edge. The flux of so-called gyrating ions show a clear maximum in association with SLAMS. This indicates that SLAMS indeed play an important role for pre-dissipation of the solar wind upstream of the shock.

This thesis describes and evaluates the design of nanospacecraft based on advanced multifunctional microsystems building blocks. These systems bring substantial improvements of the performance of nanosatellites and enable new space exploration, e.g. interplanetary science missions using minute space probes. Microsystems, or microelectromechanical systems, allows for extreme miniaturization using heritage from IC industry. Reducing mass and volume of spacecraft gives large savings in terms of launch costs. Definition and categorization of system and module level features in multifunctional microsystems are used to derive a spacecraft optimization algorithm which is compatible with commonly used concurrent engineering methods. The miniaturization of modules enables modular spacecraft architectures comprising powerful multifunctional microsystems, which are applicable to satellites between 10 and 1000’s of kg. This kind of complete spacecraft architecture has been developed for the NanoSpace-1 technology demonstrator satellite. The spacecraft bus uses multifunctional design to enable distributed intelligence and autonomy, graceful degradation, functional surfaces, and distributed power systems. The increase in performance of the new spacecraft architecture as compared with conventional nanosatellites is orders of magnitudes in terms of power storage, scientific payload mass ratio, pointing stabilization, and long time space operation. This high-performance system-of-microsystems architecture has been successfully employed on two space robotic concepts: a miniaturized submersible vehicle for Jupiter’s Moon Europa and a miniaturized spherical robot. The submersible is enabled by miniaturization of electronics into 3-dimensional, vertically integrated multi-chip-modules together with new interconnection methods. These technologies enabled the submersible vehicle tube-shaped design within 20 cm length and 5 cm diameter. The spherical rover was developed for long range and networked science investigations of interplanetary bodies. The rover weighs 3.5 kg and is shown to endure direct reentry on Mars, which increases the ratio between the landed mobile payload mass and the initial mass in Mars orbit by a factor of 18.

QC 20100901

Estimation of the total mission life of a spacecraft is an important issue for the communication satellite industries. For accurate determination of the remaining mission life of the satellite it is essential to estimate the amount of propellant present in the propellant tank of the spacecraft at various stages of its mission life. Because the annual revenue incurred from a typical communication satellite operating at its full capacity is on the order of millions of dollars, premature removal of spacecraft from their orbits results in heavy losses. Various techniques such as the bo okkeeping method, the gas law method, numerical modeling techniques, and use of capacitive sensors have been employed in the past for accurate determination of the amount of propellant present in a spacecraft. First half of the thesis is concerned with sensitivity analysis of the various propellant gauging techniques, that is, estimating the e ects of the uncertainty in the instruments employed in the propellant gauging system on the onboard propellant estimation. This sensitivity analysis is done for three existing propellant gauging techniques – gas injection method, book-keeping method and the propellant tank heating method. A comparative study of the precision with which the onboard propellant is estimated by the three techniques is done and the primary source of uncertainty for all the three techniques is identified. It is illustrated that all the three methods — the gas injection method, the book-keeping method and the propellant tank heating method — are inherently indirect methods of propellant gauging, as a consequence of which, the precision with which the three techniques estimate the residual propellant decreases towards the end of mission life of the spacecraft. The second half of the thesis explores the possibility of using a new propellant tank configuration, consisting of a truncated cone centrally mounted within a spherical propellant tank, to measure the amount of liquid propellant present within the tank. The liquid propellant present within the propellant tank orients itself in a geometry, by virtue of its dominant surface tension force in zero-g condition, which minimizes its total surface energy. Study reveals that the amount of liquid propellant present in the tank can thus be estimated by measuring the height of the propellant meniscus within the central cone. It is also observed that, unlike gas law metho d, bookkeeping method or the propellant tank heating metho d, where the precision of the estimated propellant fill-fraction decreases towards the end-of-life of the spacecraft, for the proposed new configuration the precision increases.

Orbit of a spacecraft in three dimensional Inertial Reference Frame is in general represented by a standard set of six parameters like Keplerian Orbital Elements namely semimajor axis, eccentricity, inclination, argument of perigee, right ascension of ascending node, and true anomaly. An orbit can also be represented by an equivalent set of six parameters namely the position and velocity vectors, hereafter referred as orbit-vectors. The process of determining the six orbital parameters from redundant set of observations (more than the required minimum observations) is known as Orbit Determination (OD) process. This is, in general, solved using Least Squares principle. Availability of accurate, almost continuous, space borne observations provide tremendous scope for simplifications and new directions in Autonomous OD (AOD). The objective of this thesis is to develop a suitable scheme for onboard autonomy in OD, specifically for low-earth-orbit-missions that are in high demand in the immediate future. The focus is on adopting a simple orbit model by a thorough study and analysis by considering the individual contributions from the different force models or component accelerations acting on the spacecraft. Second step in this work is to address the application of an onboard estimation scheme like Kalman Filter for onboard processing. The impact of the approximation made in the orbit model for filter implementation manifests as propagation error or estimation residuals in the estimation. The normal procedure of tuning the filter is by getting an appropriate state and measurement noise covariance matrices by some means, sometimes through trial and error basis. Since this tuning is laborious and the performance may vary with different contexts, it is attempted to propose a scheme on a more general footing, with dynamically compensating for the model simplification. There are three parts of this problem namely (i) Analysis of different Orbit Dynamics Models and selection of a simplified Onboard Model (ii) Design of an Estimator Filter based on Kalman Filter approach for Onboard Applications and (iii) Development of a suitable Filter Compensation procedure to ensure best estimates of orbit vectors even with the simplified orbit model. Development of a Numerical Integration scheme (and a software tool) and extensive simulation exercises to justify the conclusion on the simple model to be used in the estimation procedure forms the first part of the thesis. Tables quantify the effect of individual accelerations and demonstrate the effects of various model components on orbit propagation. In general, it is well known that the atmospheric drag is a non-conservative force and reduces energy; it is also known that the effect of first zonal harmonic term is predominant than any other gravity parameters; such anticipated trends in the accuracies are obtained. This particular exercise is carried out for orbits of different altitudes and different inclinations. The analysis facilitates conclusions on a limited model orbit dynamics suitable for onboard OD. Procedures and results of this model selection analysis is published in Journal of Spacecraft Technology, Vol. 16, No.1,pp 8-30, Jan 2006, titled “Orbit Model Studies for Onboard Orbit Estimation” [69]. Design of Estimator based on Kalman Filter There are two steps involved in dealing with the next part of the defined work: • Design and implementation of Extended Kalman Filter Estimation (EKF) scheme • Steps to compensate for approximation made in the reduced orbit dynamics The GPS receivers on board some of the IRS satellites (for example, the Resource-Sat-1), output the GPS Navigation Solutions (GPSNS) namely the position and velocity vectors of the IRS satellite along with the Pseudo-range measurements. These are recorded onboard for about two orbits duration, and are down loaded. An Extended Kalman Filter Algorithm for the estimation of the orbit vectors using these GPSNS observations is developed. Estimation is carried out assuming a Gaussian white noise models for the state and observation noises. The results show a strong dependence on the initial covariance of the noise involved; reconstruction of the observations results only if the assumption of realistic noise characteristics (which are unknown) is strictly adhered. Hence this simple non-adaptive EKF is found inadequate for onboard OD scheme. Development of the Dynamics Filter Compensation (DFC) Scheme In next part of the thesis, the problem of dealing with the un-modeled accelerations has been addressed. A suitable model-compensation scheme that was first developed by D.S Ingram el at [60] and successfully applied to Lunar missions, has been modified suitably to treat the problem posed by the reduced orbit dynamics. Here, the un-modeled accelerations are approximated by the OU stochastic process described as the solution of the Langavin stochastic differential equation. A filter scheme is designed where the coefficients of the un- modeled acceleration components are also estimated along with the system state yielding a better solution. Further augmentation to the filter include a standard Adaptive Measurement Noise covariance update; results are substantiated with actual data of IRS-P6 (Resource–Sat 1, see chapter 4). Classified as the Structured Adaptive Filtering Scheme, this results in a Dynamic Filter Compensation(DFC) Scheme which provides distinctly improved results in the position of the state. First, the estimation is carried out using actual GPS Navigation Solutions as observations. What is to be estimated itself is observed; the State-Observation relation is simple. The results are seen to improve the orbit position five times; bringing down the position error from 40 meters to about 8 meters. However, this scheme superimposes an extra factor of noise in the velocity vector of the GPSNS solutions. It is noted that this scheme deals only with the process noise covariance. To tackle the noise introduced in the velocity components, modifications of the original scheme by introducing an adaptive measurement noise covariance update is done. This improves the position estimate further by about 2 meters and also removes the noise introduced in the velocity components and reconstructs the orbit velocity vector output of the GPSNS. The results are confirmed using one more set of actual data corresponding to a different date. This scheme is shown to be useful for obtaining continuous output –without data gaps- of the GPSNS output. Next, the estimation is carried out taking the actual GPS observations which are the Pseudo Range, Range rate measurements from the visible GPS satellites (visible to the GPS receiver onboard ). Switching over to the required formulation for this situation in the state-measurement relation profile, estimation is carried out. The results are confirmed in this case also. Clear graphs of comparisons with definitive orbital states (considered as actual) versus estimated states show that the model reduction attempted at the first part has been successfully tackled in this method. In this era of space-borne GPS observations, where frequent sampling of the orbiting body is suggestive of reduced orbit models, an attempt for replacement of the conventional treatment of expensive and elaborate OD procedure is proved feasible in this thesis work.

Estimation of the total mission life of a spacecraft is an important issue for the communication satellite industries. For accurate determination of the remaining mission life of the satellite it is essential to estimate the amount of propellant present in the propellant tank of the spacecraft at various stages of its mission life. Because the annual revenue incurred from a typical commu-nication satellite operating at its full capacity is on the order of millions of dollars, premature removal of spacecraft from their orbits results in heavy losses. Various techniques such as the bookkeeping method, the gas law method, numerical modeling techniques, and use of capacitive sensors have been employed in the past for accurate determination of the amount of propellant present in a spacecraft. First half of the thesis is concerned with sensitivity analysis of the various propellant gauging techniques, that is, estimating the effects of the uncertainty in the instruments employed in the propellant gauging system on the onboard propellant estimation. This sensitivity analysis is done for three existing propellant gauging techniques – gas injection method, book-keeping method and the propellant tank heating method. A comparative study of the precision with which the onboard propellant is estimated by the three techniques is done and the primary source of uncertainty for all the three techniques is identified. It is illustrated that all the three methods — the gas injection method, the book-keeping method and the propellant tank heating method — are inherently indirect methods of propellant gauging, as a consequence of which, the precision with which the three techniques estimate the residual propellant decreases towards the end of mission life of the spacecraft. The second half of the thesis explores the possibility of using a new propellant tank configuration, consisting of a truncated cone centrally mounted within a spherical propellant tank, to measure the amount of liquid propellant present within the tank. The liquid propellant present within the propellant tank orients itself in a geometry, by virtue of its dominant surface tension force in zero-g condition, which minimizes its total surface energy. Study reveals that the amount of liquid propellant present in the tank can thus be estimated by measuring the height of the propellant meniscus within the central cone. It is also observed that, unlike gas law method, bookkeeping method or the propellant tank heating method, where the precision of the estimated propellant fill-fraction decreases towards the end-of-life of the spacecraft, for the proposed new configuration the precision increases.

Coulomb forces acting between close flying charged spacecraft provide near zero propellant relative motion control, albeit with added nonlinear coupling and limited controllability. This novel concept has numerous potential applications, but also many technical challenges. In this dissertation, two- and three-craft Coulomb formations near GEO are investigated, using a rotating Hill frame dynamical model, that includes Debye shielding and differential gravity. Aspects of dynamical systems theory and optimization are applied, for insights regarding stability, and how inherent nonlinear complexities may be beneficially exploited to maintain and maneuver these electrostatic formations. Periodic relative orbits of two spacecraft, enabled by open-loop charge functions, are derived for the first time. These represent a desired extension to more substantially studied, constant charge, static Coulomb formations. An integral of motion is derived for the Hill frame model, and then applied in eliminating otherwise plausible periodic solutions. Stability of orbit families are evaluated using Floquet theory, and asymptotic stability is shown unattainable analytically. Weak stability boundary dynamics arise upon adding Coulomb forces to the relative motion problem, and therefore invariant manifolds are considered, in part, to more efficiently realize formation shape changes. A methodology to formulate and solve two-craft static Coulomb formation reconfigurations, as parameter optimization problems with minimum inertial thrust, is demonstrated. Manifolds are sought to achieve discontinuous transfers, which are then differentially corrected using charge variations and impulsive thrusting. Two nonlinear programming algorithms, gradient and stochastic, are employed as solvers and their performances are compared. Necessary and sufficient existence criteria are derived for three-craft collinear Coulomb formations, and a stability analysis is performed for the resulting discrete equilibrium cases. Each specified configuration is enabled by non-unique charge values, and so a method to compute minimum power solutions is outlined. Certain equilibrium cases are proven maintainable using only charge control, and feedback stabilized simulations demonstrate this. Practical scenarios for extending the optimal reconfiguration method are also discussed. Lastly, particular Hill frame model trajectories are integrated in an inertial frame with primary perturbations and interpolated Debye length variations. This validates qualitative stability properties, reveals particular periodic solutions to exhibit nonlinear boundedness, and illustrates higher-fidelity solution accuracies. / text