Superconducting applications in propulsion systems. Magnetic insulation for plasma propulsion devices

The purpose of this paper is to review the status of knowledge of the basic concepts needed to establish design parameters for effective magnetic insulation. The objective is to estimate the effectiveness of the magnetic field in insulating the plasma, to calculate the magnitude of the magnetic field necessary to reduce the heat transfer to the walls sufficiently enough to demonstrate the potential of magnetically driven plasma rockets

On-a-chip microdischarge thruster arrays inspired by photonic device technology for plasma television

This study shows that the practical scaling of a hollow cathode thruster device to MEMS level should be possible albeit with significant divergence from traditional design. The main divergence is the need to operate at discharge pressures between 1-3bar to maintain emitter diameter pressure products of similar values to conventional hollow cathode devices. Without operating at these pressures emitter cavity dimensions become prohibitively large for maintenance of the hollow cathode effect and without which discharge voltage would be in the hundreds of volts as with conventional microdischarge devices. In addition this requires sufficiently constrictive orifice diameters in the 10µm – 50µm range for single cathodes or <5µm larger arrays. Operation at this pressure results in very small Debye lengths (4 -5.2pm) and leads to large reductions in effective work function (0.3 – 0.43eV) via the Schottky effect. Consequently, simple work function lowering compounds such as lanthanum hexaboride (LaB6) can be used to reduce operating temperature without the significant manufacturing complexity of producing porous impregnated thermionic emitters as with macro scale hollow cathodes, while still operating <1200°C at the emitter surface. The literature shows that LaB6 can be deposited using a variety of standard microfabrication techniques

Advanced Solar-propelled Cargo Spacecraft for Mars Missions

Three concepts for an unmanned, solar powered, cargo spacecraft for Mars support missions were investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: A Solar Radiation Absorption (SRA) system, a Solar-Pumped Laser (SPL) system and a solar powered magnetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sunsynchronous Earth orbit converts solar energy to laser energy. The MPD system used indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft and the SPL powered spacecraft return to Earth for subsequent missions. The MPD propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years)

MPD thruster technology

MPD (MagnetoPlasmaDynamic) thrusters demonstrated between 2000 and 7000 seconds specific impulse at efficiencies approaching 40 percent, and were operated continuously at power levels over 500 kW. These demonstrated capabilities, combined with the simplicity and robustness of the thruster, make them attractive candidates for application to both unmanned and manned orbit raising, lunar, and planetary missions. To date, however, only a limited number of thruster configurations, propellants, and operating conditions were studied. The present status of MPD research is reviewed, including developments in the measured performance levels and electrode erosion rates. Theoretical studies of the thruster dynamics are also described. Significant progress was made in establishing empirical scaling laws, performance and lifetime limitations and in the development of numerical codes to simulate the flow field and electrode processes

Computational design of an experimental laser-powered thruster

An extensive numerical experiment, using the developed computer code, was conducted to design an optimized laser-sustained hydrogen plasma thruster. The plasma was sustained using a 30 kW CO2 laser beam operated at 10.6 micrometers focused inside the thruster. The adopted physical model considers two-dimensional compressible Navier-Stokes equations coupled with the laser power absorption process, geometric ray tracing for the laser beam, and the thermodynamically equilibrium (LTE) assumption for the plasma thermophysical and optical properties. A pressure based Navier-Stokes solver using body-fitted coordinate was used to calculate the laser-supported rocket flow which consists of both recirculating and transonic flow regions. The computer code was used to study the behavior of laser-sustained plasmas within a pipe over a wide range of forced convection and optical arrangements before it was applied to the thruster design, and these theoretical calculations agree well with existing experimental results. Several different throat size thrusters operated at 150 and 300 kPa chamber pressure were evaluated in the numerical experiment. It is found that the thruster performance (vacuum specific impulse) is highly dependent on the operating conditions, and that an adequately designed laser-supported thruster can have a specific impulse around 1500 sec. The heat loading on the wall of the calculated thrusters were also estimated, and it is comparable to heat loading on the conventional chemical rocket. It was also found that the specific impulse of the calculated thrusters can be reduced by 200 secs due to the finite chemical reaction rate

Closed-drift thruster investigations

Recent data obtained from a second generation closed-drift thruster design, employing Hall current acceleration is outlined. This type device is emphasized for electric propulsion for geocentric mission applications. Because geocentric mission profiles are best achieved with a specific impulse range of 1000 to 2000 s, closed-drift thrusters are well suited for this application, permitting time payload compromises intermediate of those possible with either electrothermal or electrostatic devices. A discussion is presented of the potential advantages of using a 1000 to 2000 s device for one way orbit raising of nonpower payloads. Because closed-drift thruster operation is not space charge limited, and requires only one power circuit for steady state operation, their application is technically advantageous. Beam, plasma and thrust characteristics are detailed for a range of operating conditions

Anomalous Diffusion at Edge and Core of a Magnetized Cold Plasma

Progress in the theory of anomalous diffusion in weakly turbulent cold magnetized plasmas is explained. Several proposed models advanced in the literature are discussed. Emphasis is put on a new proposed mechanism for anomalous diffusion transport mechanism based on the coupled action of conductive walls (excluding electrodes) bounding the plasma drain current (edge diffusion) together with the magnetic field flux "cutting" the area traced by the charged particles in their orbital motion. The same reasoning is shown to apply to the plasma core anomalous diffusion. The proposed mechanism is expected to be valid in regimes when plasma diffusion scales as Bohm diffusion and at high $B/N$, when collisions are of secondary importance.Comment: 9 pages, 4 figure

MPD thruster research issues, activities, strategies

The following activities and plans in the MPD thruster development are summarized: (1) experimental and theoretical research (magnetic nozzles at present and high power levels, MPD thrusters with applied fields extending into the thrust chamber, and improved electrode performance); and (2) tools (MACH2 code for MPD and nozzle flow calculation, laser diagnostics and spectroscopy for non-intrusive measurements of flow conditions, and extension to higher power). National strategies are also outlined

Development and Application of Multidimensional Computational Models for Hall Thrusters

The goal of this work is to aid the development of high-power Hall-effect thrusters through modeling and simulation. The focus is both on improving the state-of-the-art in the field of Hall thruster numerical simulation, as well as studying several physical processes that are important to Hall thruster development and application. Since Hall thrusters have been in use for more than half a century, they have built a reputation of reliability, however they are known for low power operation with primary applications such as station-keeping and orbit raising. Within the past decade there has been a significant effort to increase the power levels for these electric propulsion devices, but when considering such recent developments, several problems become apparent. First, as we scale these devices to higher power, higher flow rates and more propellant are needed. This translates into increased costs for ground testing, as well as in-space operation. These issues are addressed through a study of an alternative and less ex- pensive option to the ubiquitous xenon gas: krypton. This new chemical species was added to the Hall2De simulation framework and two thrusters were simulated with krypton propellant. Computed thrust values were found to be within 6% for xenon, and within the 2% experimental measurement error for krypton. Next, scaling to higher power leads to more energetic ions impacting the thruster surfaces that may in turn lead to higher observed erosion rates. Therefore, we must consider the problem of discharge channel erosion, which is investigated by simulating an optical experimental diagnostic that is meant to non-invasively determine the erosion rate: cavity-ring-down spectroscopy. The simulation result over predicts the boron number density in the plume by a factor of 3, and this may be attributed to the significant (±50%) uncertainty in the thruster operation time. Further, the desire to scale Hall thrusters to higher power has led to the idea of con- centrically nesting multiple discharge channels into a single thruster. This novel con- figuration has yielded anomalous thrust gains which have been investigated through a cold gas (neutral) simulation of dual channel operation. In conjunction with significant experimental work performed by colleagues at the Plasmadynamics and Electric Propulsion Laboratory (PEPL) it was found that the anomalous thrust gains may be explained based on the near-plume pressure distribution. In an effort to fully characterize the thruster, a plasma simulation of the single channel mode operation was performed, and thrust was matched to within 9%, while discharge current was matched to within 5% of the measured values. Moreover, it was determined that improved modeling capabilities are required in order to simulate the dual-channel or even independent outer-channel operating modes. Therefore, a new Cartesian 2D axisymmetric electron fluid model is developed, verified and then integrated within an existing state-of-the-art hybrid-particle-in-cell framework.PHDAerospace EngineeringUniversity of Michigan, Horace H. Rackham School of Graduate Studieshttps://deepblue.lib.umich.edu/bitstream/2027.42/147521/1/horatiud_1.pd

3D Simulation of Plume Flows from a Cluster of Plasma Thrusters

Peer Reviewedhttp://deepblue.lib.umich.edu/bitstream/2027.42/77194/1/AIAA-2005-4662-529.pd