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Donius B.R.,Missouri University of Science and Technology | Donius B.R.,Aerospace Plasma Laboratory | Rovey J.L.,Missouri University of Science and Technology
Journal of Spacecraft and Rockets | Year: 2011

Analytical and numerical investigations of the performance of a series of potential dual-mode propulsion systems using ionic liquids are presented. A comparison of the predicted specific impulse of ionic liquids with hydrazine and unsymmetrical dimethylhydrazine shows that ionic liquid fuels have a 3-12% lower specific impulse when paired with a nitrogen tetroxide oxidizer. However, when paired with hydroxylammonium nitrate oxidizer, the specific impulse of the ionic liquids is 1-4% lower than that of hydrazine and unsymmetrical dimethylhydrazine paired with nitrogen tetroxide. Analytical investigation of an electrospray electric propulsion system shows that ion extraction in the pure ion regime provides a very high specific impulse, outside the optimum range for potential missions. Results suggest a deceleration grid, a lower ion fraction, or emission of higher solvated states is required. Analysis of a dualmode ionic-liquid-propelled spacecraft shows that the electric propulsion component determines the overall feasibility compared with current technology. Results indicate that the specific power for an ionic liquid electrospray system must be at least 15 W/kg in order for a dual-mode ionic liquid system to compete with traditional hydrazine and Hall thruster technology. Copyright © 2010 by Joshua L. Rovey. Published by the American Institute of Aeronautics and Astronautics, Inc.


Berg S.P.,Missouri University of Science and Technology | Berg S.P.,Aerospace Plasma Laboratory | Rovey J.L.,Missouri University of Science and Technology | Rovey J.L.,Aerospace Plasma Laboratory
49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference | Year: 2013

A novel multi-mode spacecraft propulsion concept is presented. The concept combines chemical monopropellant and electric pulsed inductive thruster technology to include shared propellant and shared conical nozzle. Geometry calculations show that existing conical pulsed inductive thruster experiments are typical of large (1000-4000 N) chemical monopropellant thruster nozzles. Performance and propulsion system mass required to accomplish a 1500 m/s delta-V with a 500 kg payload was calculated for geometries including 20-55 degree divergence angles. Results show that combining nozzle geometry is not beneficial in terms of propulsion system mass for small nozzle divergence angles, however using a nozzle with a 55 degree divergence angle results in a 1-2% reduction in propulsion system mass compared to an equivalent thrust system utilizing a separate chemical bell nozzle and flat coil PIT device despite having 19% lower chemical specific impulse and 18% lower electric thrust efficiency. Results suggest that using even larger divergence angles could yield even more benefit.


Ferry J.W.,Missouri University of Science and Technology | Ferry J.W.,Aerospace Plasma Laboratory | Rovey J.L.,Missouri University of Science and Technology
5th Flow Control Conference | Year: 2010

Plasma-based aerodynamic actuators can modify a flow field without the need for moving control surfaces or a source of pressurized air. Actuator power consumption and thrust production were measured for driving frequencies between 1 and 18 kHz, and for driving voltages of 6 and 9 kV peak to peak. The actuator consumed between 3 and 22 W, and produced thrust levels between 0.05 and 0.2 mN per meter span. A comparison of results showed good agreement between this work and previous authors' results. The actuator effectiveness (thrust produced per watt of power input) was found to range between 0.017 and 0.11 mN/W. The continuous power consumption of a DBD actuator-based control system was then estimated by modeling the actuators as jet flaps. The elevator jet flap strength required to trim a small aircraft in flight was determined. A 0.5 kg aircraft with 0.76 m2 wing area required between 0.47 and 2.22 kW of power for trim. A 3 kg aircraft with 1.27 m2 wing area required between 13.6 and 54.6 kW of power for trim. In the most challenging circumstances, flight at stall or max velocity, current battery capacities would allow these aircraft to maintain trimmed flight for only 73 seconds. © 2010 by Joseph W. Ferry.


Satonik A.J.,Missouri University of Science and Technology | Satonik A.J.,Aerospace Plasma Laboratory | Rovey J.L.,Missouri University of Science and Technology | Rovey J.L.,China Aerospace Science and Technology Corporation | Hilmas G.,Missouri University of Science and Technology
Journal of Propulsion and Power | Year: 2014

Worn Hall-effect thrusters show a variety of unique microstructures and elemental compositions in the boron nitride thruster channel walls. Understanding the plasma conditions that lead to the formation of these microstructures and elemental changes can assist in the goal of creating new ceramic materials with desired plasma material interactions. Pristine and worn Hall-effect thruster channel samples of boron nitride were exposed to xenon plasma in a magnetron sputter device. Erosion rate was shown to depend on the grade of the boron nitride ceramic and the preparation of the surface before plasma exposure. Worn Hall-effect thruster thruster channel samples eroded up to 90% faster than their pristine counterparts. This result highlights the evolution and feedback of the plasma-material interaction within the Hall-effect thruster channel. Microscope images of the ceramic surface show that the magnetron plasma rounded the edges of the ceramic grains to closely match the worn Hall-effect thruster surface. This effect is absent from pure ion beam bombardment and appears to be unique to quasi-neutral plasma exposure. Copyright © 2013 by the American Institute of Aeronautics and Astronautics, Inc.


Zidar D.G.,Missouri University of Science and Technology | Zidar D.G.,Aerospace Plasma Laboratory | Rovey J.L.,Missouri University of Science and Technology
Journal of Propulsion and Power | Year: 2012

Surface properties of Hall-effect thruster channel walls play an important role in the performance and lifetime of the device. Physical models of near-wall effects are beginning to be incorporated into thruster simulations, and these models must account for the evolution of channel surface properties due to thruster operation. Results from this study show differences in boron nitride channel surface properties from beginning-of-life and after hundreds of hours of operation. Two worn thruster channels of different boron nitride grades are compared with their corresponding pristine and shadow-shielded samples. Pristine HP-grade boron nitride surface roughness is 9000 ± 700 å, whereas the worn sample is 110,900 ± 8900 å at the exit plane. Pristine M26-grade boron nitride surface roughness is 18400 ± 1400 å, whereas the worn sample is 52300 ± 4200 å at the exit plane. Comparison of pristine and worn channel surfaces also shows surface properties are dependent on the axial position within the channel. For example, surface roughness increases by as much as a factor of 5.4, and the surface atom fraction of carbon and metallic atoms decreases by a factor of 2.9 from anode to the exit plane. Macroscopic striations at the exit plane angled 10-30° from the axial are found to be related to the electron gyroradius and give rise to anisotropic surface roughness. Smoothing of ceramic grains at the microscopic level is also found. © 2011 by David G. Zidar.


Hu J.,Missouri University of Science and Technology | Hu J.,Aerospace Plasma Laboratory | Rovey J.L.,Missouri University of Science and Technology
Journal of Applied Physics | Year: 2013

In this paper, a retarding potential energy analyzer (RPEA) specific for pulsed electron beams within the pressure range of tens of mTorr is developed and used to investigate the energy of transient hollow cathode discharge produced electron beams. This RPEA has been applied in a pseudospark-based electron beam source at applied potential up to 20 kV. Experimental investigations under applied potential of 5 kV, 10 kV, 15 kV, and 20 kV were carried out and the time-resolved electron energy distributions are constructed. The numbers of electrons within various energy groups are calculated from the time-resolved electron energy spectra. Results show that the maximum number of electrons is emitted within the energy range of 40%-60% of the full applied potential on the pseudospark device, and varies from 22.5 ± 2.0% to 38.9 ± 2.0 % of the total number of emitted electrons. Additionally, the energy transformation efficiency of stored electrical energy to electron beam energy is calculated from presented data. The energy transformation efficiency increases from 11.4 ± 0.5% at 5 kV breakdown voltage to 23.2 ± 3.5% at 20 kV breakdown voltage. © 2013 AIP Publishing LLC.


Berg S.P.,Missouri University of Science and Technology | Berg S.P.,Aerospace Plasma Laboratory | Rovey J.L.,Missouri University of Science and Technology | Rovey J.L.,Aerospace Plasma Laboratory
Journal of Propulsion and Power | Year: 2013

Potential dual-mode monopropellant/electrospray-capable mixtures of hydroxylammonium nitrate with ionic liquid fuels [Bmim][NO3] and [Emim][EtSO4] are synthesized and tested for catalytic decomposition in a microreactor setup. The setup is benchmarked using a 30% hydrogen peroxide solution decomposed via silver catalyst. Results show similar trends but with variance in the quantitative data obtained in the literature. This was found to be a direct result of the sample-holder geometry. Hydrazine decomposition was conducted on an unsupported iridium catalyst. The same trends in terms of pressure-rise rate during decomposition (̃160 mbar/s) are obtained with unsupported catalyst but at 100 °C instead of room temperature for tests conducted on supported catalysts in the literature. For the [Bmim][NO 3]/hydroxylammonium nitrate propellant, rhenium catalyst preheated to 160 °C yielded a pressure-rise rate of 26 mbar/s, compared to 14 mbar/s for iridium catalyst and 12 mbar/s for no catalyst at the same temperature. [Emim][EtSO4]/hydroxylammonium nitrate propellant shows slightly less activity at 160 °C preheat temperature, yielding a pressure-rise rate of 20, 4, and 2.5 mbar/s for injection onto rhenium, iridium, and the thermal plate, respectively. Final results indicate that desirable ignition performance may potentially be obtained by using a supported rhenium catalyst.


Berg S.P.,Missouri University of Science and Technology | Berg S.P.,Aerospace Plasma Laboratory | Rovey J.L.,Missouri University of Science and Technology
Journal of Propulsion and Power | Year: 2013

Imidazole-based ionic liquids are investigated in terms of dual-mode chemical monopropellant and electrospray rocket propulsion capabilities. A literature review of ionic liquid physical properties is conducted to determine an initial, representative set of ionic liquids that shows favorable physical properties for both modes, followed by numerical and analytical performance simulations. The ionic liquids 1-butyl-3-methylimidazolium dicyanamide, 1-butyl-3-methylimidazolium nitrate, and 1-ethyl-3-methylimidazolium ethyl sulfate meet or exceed the storability properties of hydrazine, and their electrochemical properties indicate that they may be capable of electrospray emission in the purely ionic regime. These liquids are projected to have 13-23% reduced monopropellant propulsion performance in comparison to hydrazine due to the prediction of solid carbon formation in the exhaust. The use of these ionic liquids as a fuel component in a binary monopropellant mixture with hydroxylammonium nitrate shows a 1-4% improved specific impulse over some "green" monopropellants. Also, this avoids volatility issues and reduces the number of electrospray emitters by 18-27% and the power required by 9-16%, with oxidizing ionic liquid fuels providing the greatest savings. A fully oxygen-balanced ionic liquid will exceed the state-of-the-art performance in both modes but will require advances in hardware technology in both modes to achieve minimum functionality. Copyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc.

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