Thesis (MScEng (Mechanical and Mechatronic Engineering))--University of Stellenbosch, 2007. / The focus of this project was to investigate the working of a liquefied gas micro satellite thruster. An introduction is given in which the significance of the project in relation to the literature is stated. The objectives of the project are also stated. In the literature survey the historical development and design specifications of some relevant thruster systems is discussed. An experimental model was designed and built to test the working of a thruster system. Attention is also given to the measurement and calibration techniques used to obtain experimental data. A computer program was written to simulate the thruster system. The experimental set-up was designed so that an accumulator could be charged with liquid butane from a storage tank. The accumulator was charged with 13 ml of liquid butane, which was heated and then exhausted through a nozzle. Copper mesh was placed in the accumulator to improve the heat transfer to the butane vapour before it was exhausted through the nozzle. A cantilever beam was used to measure the thrust of the system. The system was tested under atmospheric conditions of 100 000 Pa as well as under vacuum conditions of 20 Pa. Two nozzles were also tested: nozzle-1 with a throat diameter of 1 mm and an exit diameter of 5 mm and nozzle-2 with a throat diameter of 1 mm and an exit diameter of 1.6 mm. A computer program was written to simulate the flow of the butane vapour through the nozzle, as well as the complex two-phase behaviour of the butane in the accumulator. Traditional gas dynamic theory was used to model the flow through the nozzle. The transient behaviour of the system was modelled to predict the rate of liquid to vapour mass transfer in the accumulator. Additionally, the computer program was developed to simulate the system with copper mesh placed in the accumulator. From the experimental results it was shown that the addition of copper mesh in the accumulator improved the total thrust achieved with a 13 ml charge of liquid butane by more than 50 %. Under atmospheric conditions shockwaves were present in both of the two nozzles tested. Nozzle-2 showed an increase of 91 % in the total thrust achieved over a 5 second burst compared to the total thrust achieved using nozzle-1. With no copper mesh in the accumulator and using nozzle-1 a peak thrust of 39 mN was achieved under atmospheric conditions while under vacuum conditions a peak thrust of 495 mN was achieved. This resulted in a total thrust of 0.365 Ns under atmospheric conditions and 4.88 Ns under vacuum conditions with a 13 ml charge of liquid butane. Using the total thrust achieved the specific impulse of the system was calculated as 5 seconds under atmospheric conditions and 67.5 seconds under vacuum conditions with no mesh in the accumulator and using nozzle-1. The theoretical model compared well with the experimental results except when nozzle-1 was modelled under atmospheric conditions. Under vacuum conditions the results obtained from the theoretical model compared well with the experimental results using both of the nozzles. In the modelling of the mesh in the accumulator an overall heat transfer factor was incorporated into the model to take into account the uncertainty of the heat transfer area as well as the overall heat transfer coefficient. The theoretical model and experimental test results are discussed and thereafter conclusions are also drawn. There are also recommendations made for future work that could be done in the further development of a liquefied gas micro satellite thruster system. It is recommended that a “resistojet” type thruster should be tested at the University of Stellenbosch and that further testing be done with mesh in the accumulator to find the optimum amount of mesh that should be placed in the accumulator.

Thesis (MScEng)--University of Stellenbosch, 2003. / ENGLISH ABSTRACT: The focus of this thesis is on the investigation of the functioning of a liquefied-gas thruster. Such a thruster could be used to provide secondary propulsion to a microsatellite in orbit. A general overview of the need for thrusters in micro-satellites is put forward in the introduction. Motivation for deciding to investigate a liquefied-gas system is presented. Recent developments in the field of micro-satellites are discussed as well as their relevance to the project undertaken. Fundamental background theory relevant to the engineering problems associated with the development and analysis of such a system is also presented. Computer programs were written to simulate such a liquefied-gas thruster system. The experimental work carried out to analyse the system from a practical view-point is documented. Attention is also given to the measurement and calibration techniques used to obtain experimental data. One-dimensional fully explicit transient mathematical models of the thruster system were developed to model the system using both compressed air and butane as propellants. These models were incorporated into computer programs used to simulate the transient behaviour of the system. Although it is intended to use butane as the propellant onboard a satellite, the reason for modelling and simulating a system using compressed air is because air is a convenient fluid to work with from both a theoretical and practical point of view. An experimental model of a thruster system was designed, built and tested using air and butane as propellants. Most of the model was built using perspex to allow for the observation of the two-phase behaviour of the propellant inside the system. Locally purchased components were used for the solenoid and fill valves. Readily available butane lighter fluid was used for butane testing. Self-made heating elements were used to provide heat input to the propellant. Testing was done at different back pressures ranging from 100 kPa down to 20 kPa in a vacuum chamber. Good comparison between theoretical and experimental results was obtained for air. Theoretical results for peak thrusts tended to over predict experimental results by approximately 15 % for a system exhausting to a pressure of 100 kPa. Peak thrusts as high as 0.2 N were obtained for vacuum tests conducted at an absolute pressure of 20 kPa. Peak thrusts of approximately 50 mN were achieved for experimental testing III atmospheric conditions using butane with a starting pressure of between 270 and 290 kPa. Typical average thrusts of between 20 mN and 30 mN were noted for butane testing with initial pressure of between 200 to 300 kPa. Peak thrusts of over 0.1 N were observed for vacuum testing at an absolute pressure of 20 kPa. An equation to correlate the experimentally determined average thrust as a function of pulse duration and starting pressure was developed. This correlated most of the experimental data to within ±25 %. Theoretical results for butane testing are able to predict peak thrusts within approximately 20 % for starting pressures in the range of 200 to 300 kPa. Since the project was an exploratory investigation into a liquefied-gas thruster, some additional aspects relating to such systems were also given attention. The effect of liquid propellant motion or sloshing was considered and recommendations regarding the design and placement of the propellant tanks were made. The use of heat pipes as an alternative to electrical heating elements was investigated and some elementary design aspects are presented graphically. The management of the liquid propellant using surface tension devices was examined qualitatively. Recommendations relating to future projects in the field of simple, low-cost propulsion systems for micro-satellites are put forward. More specifically these recommendations are with regard to: thermo-fluid modelling of the propellant, future experimental work to be done, techniques to measure small thrusts and vacuum chamber testing. / AFRIKAANSE OPSOMMING: Die tesis ondersoek die funksionering van 'n vervloeidegas stuwer. So 'n stuwer kan gebruik word om sekondêre aandrywing aan 'n mikro-satelliet in 'n wentelbaan te verskaf. 'n Algemene oorsig oor die behoeftes van stuwers vir mikro-satelliete word voortgesit in die inleiding. Redes vir die gebruik van 'n vervloeidegas stuwer word bespreek. Onlangse ontwikkelinge in die veld van mikro-satelliet aandrywing word bespreek asook die toepaslikheid daarvan. Fundamentele teoretiese agtergrond verbonde aan die ontwikkeling en analise van so 'n stuwer stelsel word ook gegee. Rekenaarprogramme is geskryf om die gedrag van so 'n stuwer stelsel te simuleer. Eksperimentele werk is gedoen om die stelsel vanuit 'n praktiese oogpunt te analiseer. Aandag word ook gegee aan die metings- en kalibrasietegnieke soos toegepas vir die eksperimentele werk. Eendimensionele volle eksplisiete wiskundige modelle is ontwikkelom die oorgangsgedrag van die stuwer-stelsel te simuleer met beide lug en butaan as dryfmiddel. Hierdie modelle is geïnkorporeer in die rekenaar programme om die stuwer stelsel te simuleer. Alhoewel dit beoog word om butaan as die dryfmiddel aan boord die satelliet te gebruik, is lug ook gebruik vir simulasie weens sy gerieflikheid as 'n vloeier uit beide 'n teoretiese en 'n praktiese oogpunt. 'n Eksperimentele model van die stuwer stelsel is ontwerp, gebou en getoets met beide lug en butaan as dryfmiddels. Die model is hoofsaaklik uit perspex gebou sodat die twee-fase gedrag van die butaan uitgebeeld kon word. Vrylik beskikbare butaan aansteker vloeistof IS gebruik VIr butaan toetsing. Selfvervaardigde verhittingselemente is gebruik om hitte aan die dryfmiddel te verskaf. Toetse is gedoen deur verskeie omgewingsdrukke varieërend van 100 kPa af tot 20 kPa in 'n vakuumtenk te gebruik. Goeie ooreenstemming tussen die teoretiese en eksperimentele resultate vir die toetsing van lug is verkry. Die teoretiese resultate neig om die piek stukrag 15 % hoër te voorspel as die eksperimentele resultate vir 'n stelsel wat tot 'n omgewingsdruk van 100 kPa by die uitlaat. Piek stukragte van meer as 0.2 N is gekry vir vakuum toetse wat gedoen is by 'n omgewingsdruk van 20 kPa. Tydens eksperimentele toetsing met butaan teen 'n aanvanklike druk tussen 270 en 290 kPa, in atmosferiese toestande, is piek stukragte van ongeveer 50 mN behaal. Tipiese gemiddelde stukragte van tussen 20 en 30 mN is waargeneem vir butaan toetsing teen 'n aanvanklike druk tussen 200 en 300 kPa. Piek stukragte van meer as 0.1 N is behaal vir vakuum toetse met 'n absolute druk van 20 kPa. 'n Vergelyking om die gemiddelde stukrag, wat eksperimenteel bepaal is, as 'n funksie van puls tydsduur en aanvanklike druk te korreleer, is ontwikkel. Die meeste eksperimentele data se afwyking van die korrelasie-vergelyking was minder as 25 %. Teoretiese resultate vir butaantoetse het piek stukragte binne 20 % van die eksperimenteel metings korrek voorspel vir aanvanklike drukke tussen 200 tot 300 kPa. Weens die feit dat die projek 'n oorhoofse ondersoek in In vervloeidegas stuwer behels, is aandag ook gegee aan addisionele aspekte wat verband hou met sulke stelsels. Die effek van die vloeistof-dryfmiddel se onstabiele beweging in sy tenke is in ag geneem en voorstelle vir die ontwerp en plasing van die dryfmiddel tenke is gemaak. Die gebruik van hitte pype as 'n alternatief vir elektriese verhittingselemente is ondersoek. Verskeie ontwerp aspekte word grafies voorgestel. Die bestuur van die vloeistof-dryfmiddel deur van oppervlak spannings apparaat gebruik te maak, is kwalitatief ondersoek. Voorstelle vir verdere navorsing in die veld van eenvoudige, lae-koste stuwer stelsels vir mikro-satelliete is gemaak. Meer spesifiek is hierdie voorstelle gerig op die termo-vloeidinamiese modellering van die dryfmiddel, verdere eksperimentele navorsing, tegnieke om klein stukragte te meet en vakuumtenk toetse.

This thesis explores how an Asteroid Sample Return Mission might make use of solar electric propulsion to send a spacecraft on a journey to the asteroid 1989ML and back. It examines different trajectories that can be used to get an asteroid sample return or similar spacecraft to an interplanetary destination and back in the most fuel-efficient manner. While current plans call for keeping such a spacecraft on the asteroid performing science experiments for approximately 90 days, it is prudent to inquire how lengthening or shortening this time period may affect mission fuel requirements. Using optimal control methods, various mission scenarios have been modeled and simulated. The results suggest that the amount of time that the spacecraft may spend on the asteroid surface can be approximated as a linear function of the available fuel mass. Furthermore, It can be shown that as maximum available thrust is decreased, the radial component of the optimal thrust vector becomes more pronounced.

CIVINS / The United States Navy has recognized the rising age of its nuclear reactors. With this increasing age comes increasing importance of backup generators. In addition to the need for decay heat removal common to all (naval and commercial) nuclear reactors, naval vessels with nuclear reactors also require a backup means of propulsion. All underway Navy nuclear reactors are operated with diesel generators as a backup power system, able to provide emergency electric power for reactor decay heat removal as well as enough electric power to supply an emergency propulsion mechanism. While all commercial nuclear reactors are required to incorporate muhiple backup generators, naval submarine nuclear plants feature a single backup generator. The increasing age of naval nuclear reactors, coupled with the dual reqmrements of a submarine's solitary backup generator, makes the study of submarine backup generators vital. / CIVINS / US Navy (USN) author

The mechanical pitching characteristic of the NPS flapping-wing Micro Air Vehicle (MAV) developed by Professor Kevin D. Jones are studied experimentally through the use of constant temperature anemometry and force balance techniques. The MAV without the main fixed-wing is placed in a laminar flow field within a low speed wind tunnel with the wake after the flapping wings characterized with a constant temperature anemometer and thrust generation measured by a load cell at various neutral angles, flapping frequencies and free stream velocities. The experiments seek to determine the effects on the MAV propulsion when the neutral angle of attack of the flapping wings is varied. Flow visualization is also performed to better enhance understanding of the flow field across the pitched flapping wings.

Approved for public release, distribution is unlimited / This thesis explores using a direct pseudospectral method for the solution of optimal control problems with mixed dynamics. An easy to use MATLAB optimization package known as DIDO is used to obtain the solutions. The modeling of both low thrust interplanetary trajectories as well as aerocapture trajectories is detailed and the solutions for low thrust minimum time and minimum fuel trajectories are explored with particular emphasis on verification of the optimality of the obtained solution. Optimal aerocpature trajectories are solved for rotating atmospheres over a range of arrival Vinfinities. Solutions are obtained using various performance indexes including minimum fuel, minimum heat load, and minimum total aerocapture mass. Finally, the problem formulation and solutions for the mixed dynamic problem of low thrust trajectories with a terminal aerocapture maneuver is addressed yielding new trajectories maximizing the total scientific mass at arrival. / Lieutenant, United States Navy

A lithium vaporizer for a high-power magnetoplasmadynamic (MPD) thruster is modeled using a parallel approach. A one-dimensional, thermal-resistive network is developed and used to calculate the required vaporizer length and power as a function of various parameters. After comparing results predicted by this network model with preheat power data for a 200 kW MPD thruster, we investigate performance over a parameter space of interest for the Advanced Lithium-Fed, Applied-field, Lorentz Force Accelerator (ALFA2) thruster. Heater power sensitivity to cathode tube emissivity, mass flow rate, and vapor superheat are presented. The cold-start heater power for 80 mg/s is found to range from 3.38 to 3.60 kW, corresponding to a vaporizer (axial) length of 18 to 26 cm. The strongest drivers of vaporizer performance are cathode tube emissivity and a conduction heat sink through the mounting flange. Also, for the baseline ALFA2 case, it is shown that increasing the vapor superheat from 100 K to 300 K has the effect of lowering the vaporizer thermal efficiency from 57% to 49%. Also, a finite-volume computational fluid dynamic (CFD) is implemented in FLUENT 6.2 which includes conjugated heat transfer to the solid components of the cathode. This model uses a single-fluid mixture model to simulate the effects of the two-phase vaporizer flow with source terms that model the vaporization. This model provides a solution of higher fidelity by including three-dimensional fluid dynamics such as thermal and momentum boundary layers, as well as calculating a higher resolution temperature distribution throughout the cathode assembly. Results from this model are presented for three mass flow rates of interest (40 mg/s, 80 mg/s, and 120 mg/s). Using a fixed power and length taken from the conceptual ALFA2 design, the dryout point ranges from 12.3-17.6 cm from the base of the cathode assembly for 40 mg/s and 80 mg/s, respectively. For the 120 mg/s case, the two-phase flow never reaches dryout. Finally, results two modeling approaches are compare favorably, with a maximum disagreement of 13.0 percent in prediction of the dryout point and 4.2 percent in predictions of the exit temperature.

"A study is presented detailing the simulation of a drag-free follow-on mission to NASA’s Gravity Recovery and Climate Experiment (GRACE). This work evaluates controller performance, as well as thrust, power, and propellant mass requirements for drag-free spacecraft operation at orbital altitudes of 160 - 225 kilometers. In addition, sensitivities to thermospheric wind, GPS signal accuracy and availability of ephemeris data are studied. Orbital dynamics were modeled in Matlab and take into account 2 body gravity effects, J2-J6 non-spherical Earth effects, atmospheric drag and control thrust. A drag model is used in which the drag acceleration is a function of the spacecraft’s relative velocity to the atmosphere, and a “drag parameter,” which includes the spacecraft’s drag coefficient and local mass density of the atmosphere. A MSISE-90 atmospheric model is used to provide local mass densities as well as free stream flow conditions for a Direct Simulation Monte Carlo drag analysis used to validate the spacecraft drag coefficient. The controller is designed around an onboard inertial sensor which uses a freely floating reference mass to measure deviations in the spacecraft position, resulting from non-gravitational forces, from a desired target orbit. Thruster (control actuator) models are based on two different Hall thrusters for providing the orbital along-track acceleration, colloid thrusters for the normal acceleration, and a miniature xenon ion thruster (MiXI) for the cross-track acceleration. The most demanding propulsion requirements correspond to the lowest altitude considered, 160 kilometers. At this altitude the maximum along-track thrust component is calculated to be 98 millinewtons with a required dynamic (throttling) response of 41 mN/s. The maximum position error at this altitude was shown to be in the along-track direction with a magnitude of 3314.9 nanometers and a peak spectral content of 1800 nm/sqrt(Hz) at about 0.1 Hz. At 225 kilometers, the maximum along-track thrust component reduces to 10.3 millinewtons. The maximum dynamic response at this altitude is 4.23 mN/s. The maximum along-track position error is reduced to 367.9 nanometers with a spectral content peak of 40 nm/sqrt(Hz) at 0.1 Hz. For all altitudes, the maximum state errors increase as the mission length increases, however, higher altitude missions show less of a maximum displacement error increase over time than those of lower orbits. The ability of a colloid thruster to control the normal drift is found to be dependent on how frequently the spacecraft state data is updated. Reducing the period between updates from 10 seconds to 1 second reduces the maximum normal state error component from 199 nanometers to less than 32 nanometers, suggesting that spacecraft state update frequency could be a major driver in keeping the spacecraft on the target trajectory. Sensitivity of maximum required thrust and accumulated sensor error to measurement uncertainty is found to be less of a driver than state update frequency. A ‘worst case” thermospheric wind gust was modeled to show the increase on propulsion requirements if such an event were to occur. At 200 kilometers, maximum winds have been measured to be in increase of 650 m/s in the westward direction in the southern pole region. Assuming the majority of the 650 m/s gust occurs over a 4 second time span, the maximum required cross-track thrust at 200 kilometers increases from 1.12 to 2.01 millinewtons. This large increase may drive the thruster choice for a drag-free mission at a similar altitude. For the spacecraft point design considered with a propellant mass fraction of 0.18, the mission lifetime for the 160 km case was calculated to be 0.76 years. This increases 2.27 years at an altitude of 225 km."

"As new, smaller satellites are built, the need for improved on-board propulsion systems has grown. The pulsed plasma thruster has received attention due to its low power requirements, its simple propellant management, and the success of initial flight tests. Successful integration of PPTs on spacecraft requires the comprehensive evaluation of possible plume-spacecraft interactions. The PPT plume consists of neutrals and ions from the decomposition of the Teflon propellant, material from electrode erosion, as well as electromagnetic fields and optical emissions. To investigate the PPT plume, an on-going program is underway at WPI that combines experimental and computational investigations. Experimental investigation of the PPT plume is challenging due to the unsteady, pulsed as well as the partially ionized character of the plume. In this thesis, a triple Langmuir probe apparatus was designed and used to obtain electron temperature and density measurements in the plume of a PPT. This experimental investigation provides further characterization of the plume, much needed validation data for computational models, and is useful in thruster optimization studies. The pulsed plasma thruster used in this study is a rectangular geometry laboratory model built at NASA Lewis Research Center for component lifetime tests and plume studies. It is almost identical in size and performance to the LES 8/9 thruster, ablating 26.6 ug of Teflon, producing an impulse bit of 256 uN-s and a specific impulse of 986 s at 20 J. All experiments were carried out at NASA LeRC Electric Propulsion Laboratory. The experimental setup included triple Langmuir probes mounted on a moveable probe stand, to collect data over a wide range of locations and operating conditions. Triple probes have the ability to instantaneously measure electron temperature and density, and have the benefit of being relatively simple to use, compared to other methods used to measure these same properties. The implementation of this measuring technique is discussed in detail, to aid future work that utilizes these devices. Electron temperature and density was measured from up to 45 degrees from the centerline on planes parallel and perpendicular to the thruster electrodes, for thruster energy levels of 5, 20 and 40 J. Radial distances extend from 6 to 20 cm downstream from the Teflon surface. These locations cover the core of the PPT plume, over a range of energy levels that corresponds to proposed mission operating conditions. Data analysis shows the spatial and temporal variation of the plume. Maximum electron density near the exit of the thruster is 1.6 x 1020, 1.6 x 1021, and 1.8 x 1021 m-3 for the 5, 20 and 40 J discharges, respectively. At 20 cm downstream from the Teflon surface, densities are 1 x 1019, 1.5 x 1020 and 4.2 x 1020 for the 5, 20 and 40 J discharges, respectively. The average electron temperature at maximum density was found to vary between 3.75 and 4.0 eV for the above density measurements at the thruster exit, and 20 cm from the Teflon surface the temperatures are 0.5, 2.5, and 3 eV for the 5, 20 and 40 J discharges. Plume properties show a great degree of angular variation in the perpendicular plane and very little in the parallel plane, most likely due to the rectangular geometry of the PPT electrodes. Simultaneous electron temperature and density traces for a single thruster discharge show that the hottest electrons populate the leading edge of the plume. Analysis between pulses shows a 50% variation in density and a 25% variation in electron temperature. Error analysis estimates that maximum uncertainty in the temperature measurements to be approximately +/- 0.75 eV due to noise smoothing, and the maximum uncertainty in electron density to be +/- 60%, due to assumptions related to the triple probe theory. In addition, analysis of previously observed slow and fast ion components in the PPT plume was performed. The analysis shows that there is approximately a 3 us difference in creation time between the fast and slow ions, and that this correlates almost exactly with the half period of the oscillations in the thruster discharge current."

The capacity of humankind to mimic fish-like locomotion for engineering applications depends mainly on the availability of suitable actuators. Researchers have recently focused on developing robotic fish using smart materials, particularly Ionic Polymer-Metal Composites (IPMCs), as a compliant, noise-free, and scalable alternative to conventional motor-based propulsion systems. In this thesis, we investigate fish-like self propulsion using flexible bimorphs made of Macro-Fiber Composite (MFC) piezoelectric laminates. Similar to IPMCs, MFCs also exhibit high efficiency in size, energy consumption, and noise reduction. In addition, MFCs offer large dynamic forces in bending actuation, strong electromechanical coupling as well as both low-frequency and high-frequency performance capabilities. The experimental component of the presented work focuses on the characterization of an MFC bimorph propulsor for thrust generation in a quiescent fluid as well as the development of a preliminary robotic fish prototype incorporating a microcontroller and a printed-circuit-board (PCB) amplifier to generate high actuation voltage for battery-powered free locomotion. From the theoretical standpoint, a reliable modeling framework that couples the actuator dynamics, hydroelasticity, and fish locomotion theory is essential to both design and control of robotic fish. Therefore, a distributed-parameter electroelastic model with fluid effects and actuator dynamics is coupled with the elongated body theory. Both in-air and underwater experiments are performed to verify the incorporation of hydrodynamic effects in the linear actuation regime. For electroelastically nonlinear actuation levels, experimentally obtained underwater vibration response is coupled with the elongated body theory to predict the thrust output. Experiments are conducted to validate the electrohydroelastic modeling approach employed in this work and to characterize the performance of an MFC bimorph propulsor. Finally, a wireless battery-powered preliminary robotic fish prototype is developed and tested in free locomotion at different frequency and voltage levels.