Ad Astra Rocket Company

Webster, TX, United States

Ad Astra Rocket Company

Webster, TX, United States
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Bering E.A.,University of Houston | Diaz F.R.C.,Ad Astra Rocket Company | Diaz F.R.C.,University of Houston | Squire J.P.,Ad Astra Rocket Company | And 10 more authors.
Physics of Plasmas | Year: 2010

The VAriable Specific Impulse Magnetoplasma Rocket (VASIMR®) is a high power electric spacecraft propulsion system, capable of Is p /thrust modulation at constant power [F. R. Chang Díaz, Proceedings of the 39th AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, 8-11 Jan. 2001]. The VASIMR® uses a helicon discharge to generate plasma. This plasma is energized by an rf booster stage that uses left hand polarized slow mode waves launched from the high field side of the ion cyclotron resonance. In the experiments reported in this paper, the booster uses 2-4 MHz waves with up to 50 kW of power. This process is similar to the ion cyclotron heating (ICH) in tokamaks, but in the VASIMR® the ions only pass through the resonance region once. The rapid absorption of ion cyclotron waves has been predicted in recent theoretical studies. These theoretical predictions have been supported with several independent measurements in this paper. The single-pass ICH produced a substantial increase in ion velocity. Pitch angle distribution studies showed that this increase took place in the resonance region where the ion cyclotron frequency was roughly equal to the frequency on the injected rf waves. Downstream of the resonance region the perpendicular velocity boost should be converted to axial flow velocity through the conservation of the first adiabatic invariant as the magnetic field decreases in the exhaust region of the VASIMR®. This paper will review all of the single-pass ICH ion acceleration data obtained using deuterium in the first VASIMR ® physics demonstrator machine, the VX-50. During these experiments, the available power to the helicon ionization stage increased from 3 to 20+ kW. The increased plasma density produced increased plasma loading of the ICH coupler. Starting with an initial demonstration of single-pass ion cyclotron acceleration, the experiments demonstrate significant improvements in coupler efficiency and in ion heating efficiency. In deuterium plasma, 80% efficient absorption of 20 kW of ICH input power was achieved. No clear evidence for power limiting instabilities in the exhaust beam has been observed. © 2010 American Institute of Physics.


Longmier B.W.,University of Michigan | Squire J.P.,Ad Astra Rocket Company | Olsen C.S.,Ad Astra Rocket Company | Cassady L.D.,Jacobs Engineering | And 7 more authors.
Journal of Propulsion and Power | Year: 2014

Testing of the Variable Specific Impulse Magnetoplasma Rocket VX-200 engine was performed over a wide throttle range in a 150 m3 vacuum chamber with sufficient pumping to permit exhaust plume measurements at argon background pressures less than 1 × 10-3 Pa (1 × 10-5 torr) during firings, ensuring charge-exchange mean free paths longer than the vacuum chamber. Measurements of plasma flux, radio frequency power, propellant flow rate, and ion kinetic energy were used to determine the ionization cost of argon and krypton in the helicon discharge. Experimental data on ionization cost, ion fraction, exhaust plume expansion angle, thruster efficiency, and thrust are presented that characterize the VX-200 engine performance over a throttling range from 15 to 200 kW radio frequency power. A semiempirical model of the thruster efficiency as a function of specific impulse indicates an ion cyclotron heating efficiency of 85 ± 7%. Operation at a total radio frequency coupled power level of 200 kW yields a thruster efficiency of 72 ± 6% at a specific impulse of 4900 ± 300 s with argon propellant. A high thrust-to-power operating mode was characterized over a wide parameter space with a maximum thrust-to-power ratio of 51 ± 5 mN/kW at a specific impulse of 1660 ± 100 s for a ratio of ion cyclotron heating radio frequency power to helicon radio frequency power of 0.7:1. © 2012 AIAA.


Olsen C.S.,Ad Astra Rocket Company | Ballenger M.G.,Ad Astra Rocket Company | Carter M.D.,Ad Astra Rocket Company | Diaz F.R.C.,Ad Astra Rocket Company | And 7 more authors.
IEEE Transactions on Plasma Science | Year: 2015

Understanding the physics involved in plasma detachment from magnetic nozzles is well theorized, but lacking in large scale experimental support. We have undertaken an experiment using the 150-m3 variable specific impulse magnetoplasma rocket test facility and VX-200 thruster seeking evidence that detachment occurs and an understanding of the physical processes involved. It was found that the plasma jet in this experiment does indeed detach from the applied magnetic nozzle (peak field ∼2 T) in a two part process. The first part involves the ions beginning to deviate from the nozzle field 0.8-m downstream of the nozzle throat. This separation location is consistent with a loss of adiabaticity where the ratio of the ion Larmor radius to the magnetic field scale length (rLi|∇ B|/B) becomes of order unity and conservation of the magnetic moment breaks down. Downstream of this separation region, the dynamics of the unmagnetized ions and magnetized electrons, along with the ion momentum, affect the plume trajectory. The second part of the process involves the formation of plasma turbulence in the form of high-frequency electric fields. The ion and electron responses to these electric fields depend upon ion momentum, magnetic field line curvature, magnetic field strength, angle between the particle trajectories, and the effective momentum transfer time. In stronger magnetic field regions of the nozzle, the detached ion trajectories are affected such that the unmagnetized ions begin to flare radially outward. Further downstream as the magnetic field weakens, for higher ion momentum and along the edge of the plume, the fluctuating electric field enables anomalous cross-field electron transport to become more dominant. This cross-field transport occurs until the electric fields dissipate ∼2-m downstream of the nozzle throat and the ion trajectories become ballistic. This transition to ballistic flow correlates well with the sub-to-super Alfvénic flow transition (βk). There was no significant change observed to the applied magnetic field. © 2014 IEEE.


De Faoite D.,University College Dublin | Browne D.J.,University College Dublin | Del Valle Gamboa J.I.,Ad Astra Rocket Company | Stanton K.T.,University College Dublin
Applied Thermal Engineering | Year: 2016

Potential thermal management strategies for the plasma generation section of a VASIMR® high-power electric propulsion space thruster are assessed. The plasma is generated in a discharge tube using helicon waves. The plasma generation process causes a significant thermal load on the plasma discharge tube and on neighbouring components, caused by cross-field particle diffusion and UV radiation. Four potential cooling system design strategies are assessed to deal with this thermal load. Four polycrystalline ceramics are evaluated for use as the plasma discharge tube material: alumina, aluminium nitride, beryllia, and silicon nitride. A finite element analysis (FEA) method was used to model the steady-state temperature and stress fields resulting from the plasma heat flux. Of the four materials assessed, aluminium nitride would result in the lowest plasma discharge tube temperatures and stresses. It was found that a design consisting of a monolithic ceramic plasma containment tube fabricated from aluminium nitride would be capable of operating up to a power level of at least 250 kW. © 2015 Elsevier Ltd. All rights reserved.


De Faoite D.,University College Dublin | Browne D.J.,University College Dublin | Del Valle Gamboa J.I.,Ad Astra Rocket Company | Stanton K.T.,University College Dublin
International Journal of Heat and Mass Transfer | Year: 2014

The heat flux incident upon the inner surface of a plasma discharge tube during a helicon plasma discharge was estimated using an inverse method. Temperature readings were taken from the outer surface of the tube using thermocouples, and the temperature data were interpolated over the tube surface. A numerical inverse procedure based on the Alifanov iterative regularisation method was used to reconstruct the heat flux on the tube inner surface as a function of space and time. In contrast to previously-used inverse models for this application, the current model implements a thermal radiation boundary condition to realistically model the energy exchange in the device. Additionally in these experiments, steady-state operation was reached, and the accurate modelling of the steady-state condition was facilitated by the thermal radiation boundary condition. The variation of heat flux with helicon discharge power, propellant flowrate, and electromagnet current was studied, and it was found that the waste heat flux increased with applied RF power and propellant flowrate, and decreased with current supplied to the electromagnets, over the range of parameter variation tested. © 2014 Elsevier Ltd. All rights reserved.


De Faoite D.,University College Dublin | Browne D.J.,University College Dublin | Chang-Diaz F.R.,Ad Astra Rocket Company | Stanton K.T.,University College Dublin
Journal of Materials Science | Year: 2012

The current review uses the material requirements of a new space propulsion device, the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) as a basis for presenting the temperature-dependent properties of a range of dielectric ceramics, but data presented could be used in the engineering design of any ceramic component with complementary material requirements. A material is required for the gas containment tube (GCT) of VASIMR® to allow it to operate at higher power levels. The GCT's operating conditions place severe constraints on the choice of material. An electrically-insulating material is required with a high-thermal conductivity, lowdielectric loss factor, and high-thermal shock resistance. There is a lack of a representative set of temperaturedependent material property data for materials considered for this application and these are required for accurate thermo-structural modelling. This modelling would facilitate the selection of an optimum material for this component. The goal of this article is to determine the best material property data values for use in the materials selection and design of such components. A review of both experimentally and theoretically determined temperaturedependent and room temperature properties of several materials has been undertaken. Data extracted are presented by property. Properties reviewed are density, Young's, bulk and shear moduli, Poisson's ratio, tensile, flexural and compressive strength, thermal conductivity, specific heat capacity, thermal expansion coefficient, and the factors affecting maximum service temperature. Materials reviewed are alumina, aluminium nitride, beryllia, fused quartz, sialon, and silicon nitride. © Springer Science+Business Media, LLC 2012.


Dankanich J.W.,Gray Research Inc. | Dankanich J.W.,NASA | Vondra B.,Ad Astra Rocket Company | Ilin A.V.,Ad Astra Rocket Company
46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit | Year: 2010

The use of electric propulsion for Mars has been explored since the 1970's when men looked to travel beyond the moon. The use of electric propulsion has been recommended in several studies as a low-risk, lower cost approach to the robotic Mars sample return missions. Electric propulsion has been evaluated for delivery of Mars cargo using power systems order of magnitude beyond state-of-the-art. Electric propulsion has also been considered for fast transits to Mars supporting manned exploration activities. Results of generalized electric propulsion transits form Earth to Mars are presented. Trades are presented as a generalized assessment based on spacecraft mass-to-power ratio, trip time, and propulsion system performance including variable and constant specific impulse, and efficiency. Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc.


Davis C.,ElectroDynamic Applications, Inc. | Gilchrist B.,University of Michigan | Squire J.,Ad Astra Rocket Company
Journal of Propulsion and Power | Year: 2011

This paper describes the experimental characterization of single-pass ion cyclotron resonance heating as applied to acceleration of ions for electric propulsion. A millimeter-wave interferometer system has shown to be a clear and simple method of quantifying ionacceleration due to ion cyclotron resonance heating. The experimental work was done on the VX-10 experiment of the variable specific impulse magnetoplasma rocket concept. The perpendicular velocity of the ions generated by ion cyclotron resonance heating was converted into axial velocity by the decreasing gradient of the axial magnetic field at the exhaust of the propulsion system from conservation of the magnet moment. This increase in axial velocity is predicted to cause a decrease in density due to conservation of current in the plasma. Interferometer density measurements were taken at three different locations on the VX-10 experiment upstream and downstream of the ion acceleration zone. A clear measurement of a 25% density drop for helium and a 40% density drop for deuterium was measured downstream of the ion resonance zone characteristic of ion acceleration. © 2010 bythe American Institute of Aeronautics and Astronautics, Inc.


Longmier B.W.,Ad Astra Rocket Company | Cassady L.D.,Ad Astra Rocket Company | Ballenger M.G.,Ad Astra Rocket Company | Carter M.D.,Ad Astra Rocket Company | And 7 more authors.
Journal of Propulsion and Power | Year: 2011

The thruster efficiency and the thrust of a high power the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) prototype have been measured with the thruster installed inside a vacuum chamber with sufficient volume and pumping to simulate the vacuum conditions of space. Using an ion flux probe array and a PMFS, the exhaust of the VX-200 engine was characterized as a function of the coupled RF power and as a function of the radial and axial positions within the exhaust plume. A thruster efficiency of 56% was calculated using the force measurements and propellant flow rate with the specific impulse of 3400 s when operating at a total RF coupled power of 108 kW. The ionization cost of argon propellant was determined to be 87 eV for optimized values of RF power and propellant flow rate. Using a semiempirical model, it is predicted that the VX-200 engine will have a thruster efficiency of 61% at 4800 s when operated at 200 kW dc input power.


Longmier B.W.,Ad Astra Rocket Company | Bering III E.A.,University of Houston | Carter M.D.,Ad Astra Rocket Company | Cassady L.D.,Ad Astra Rocket Company | And 8 more authors.
Plasma Sources Science and Technology | Year: 2011

The helicon plasma stage in the Variable Specific Impulse Magnetoplasma Rocket (VASIMR®) VX-200i device was used to characterize an axial plasma potential profile within an expanding magnetic nozzle region of the laboratory based device. The ion acceleration mechanism is identified as an ambipolar electric field produced by an electron pressure gradient, resulting in a local axial ion speed of Mach 4 downstream of the magnetic nozzle. A 20 eV argon ion kinetic energy was measured in the helicon source, which had a peak magnetic field strength of 0.17 T. The helicon plasma source was operated with 25mg s-1 argon propellant and 30kW of RF power. The maximum measured values of plasma density and electron temperature within the exhaust plume were 1 × 1020 m-3 and 9eV, respectively. The measured plasma density is nearly an order of magnitude larger than previously reported steady-state helicon plasma sources. The exhaust plume also exhibits a 95% to 100% ionization fraction. The size scale and spatial location of the plasma potential structure in the expanding magnetic nozzle region appear to follow the size scale and spatial location of the expanding magnetic field. The thickness of the potential structure was found to be 104 to 105 λDe depending on the local electron temperature in the magnetic nozzle, many orders of magnitude larger than typical laboratory double layer structures. The background plasma density and neutral argon pressure were 1015 m-3 and 2 × 10-5 Torr, respectively, in a 150 m3 vacuum chamber during operation of the helicon plasma source. The agreement between the measured plasma potential and plasma potential that was calculated from an ambipolar ion acceleration analysis over the bulk of the axial distance where the potential drop was located is a strong confirmation of the ambipolar acceleration process. © 2011 IOP Publishing Ltd.

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