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NASA NTRS1986

Measurements of plasma parameters in the vicinity of the Space Shuttle

A Langmuir probe flown as part of the Plasma Diagnostics Package aboard the third Space Shuttle flight was used to determine electron densities, temperatures, and plasma potential in the vicinity of the Shuttle Orbiter. Measurements taken both in the cargo bay and 10 m above the cargo bay on the Remote Manipulator System arm are consistent with small satellite and laboratory results, in that reduced densities and elevated temperatures are observed in the Shuttle wake. The primary difference in the Shuttle measurements is one of magnitude; i.e., orders-of-magnitude density decreases and factor-of-five temperature enhancements. Analysis of data taken in (Delta N)/N turbulence can be as high as a few percent, and the most intense turbulence seems to occur near regions with a steep gradient in plasma pressure.

Murphy, G.↗

NASA NTRS2013

Hybrid-PIC Modeling of a High-Voltage, High-Specific-Impulse Hall Thruster

The primary life-limiting mechanism of Hall thrusters is the sputter erosion of the discharge channel walls by high-energy propellant ions. Because of the difficulty involved in characterizing this erosion experimentally, many past efforts have focused on numerical modeling to predict erosion rates and thruster lifespan, but those analyses were limited to Hall thrusters operating in the 200-400V discharge voltage range. Thrusters operating at higher discharge voltages (V(sub d) >= 500 V) present an erosion environment that may differ greatly from that of the lower-voltage thrusters modeled in the past. In this work, HPHall, a well-established hybrid-PIC code, is used to simulate NASA's High-Voltage Hall Accelerator (HiVHAc) at discharge voltages of 300, 400, and 500V as a first step towards modeling the discharge channel erosion. It is found that the model accurately predicts the thruster performance at all operating conditions to within 6%. The model predicts a normalized plasma potential profile that is consistent between all three operating points, with the acceleration zone appearing in the same approximate location. The expected trend of increasing electron temperature with increasing discharge voltage is observed. An analysis of the discharge current oscillations shows that the model predicts oscillations that are much greater in amplitude than those measured experimentally at all operating points, suggesting that the differences in oscillation amplitude are not strongly associated with discharge voltage.

Hall Thrusters↗

NASA NTRS2012

Design of a Laboratory Hall Thruster with Magnetically Shielded Channel Walls, Phase III: Comparison of Theory with Experiment

A proof-of-principle effort to demonstrate a technique by which erosion of the acceleration channel in Hall thrusters of the magnetic-layer type can be eliminated has been completed. The first principles of the technique, now known as "magnetic shielding," were derived based on the findings of numerical simulations in 2-D axisymmetric geometry. The simulations, in turn, guided the modification of an existing 6-kW laboratory Hall thruster. This magnetically shielded (MS) thruster was then built and tested. Because neither theory nor experiment alone can validate fully the first principles of the technique, the objective of the 2-yr effort was twofold: (1) to demonstrate in the laboratory that the erosion rates can be reduced by >order of magnitude, and (2) to demonstrate that the near-wall plasma properties can be altered according to the theoretical predictions. This paper concludes the demonstration of magnetic shielding by reporting on a wide range of comparisons between results from numerical simulations and laboratory diagnostics. Collectively, we find that the comparisons validate the theory. Near the walls of the MS thruster, theory and experiment agree: (1) the plasma potential has been sustained at values near the discharge voltage, and (2) the electron temperature has been lowered by at least 2.5-3 times compared to the unshielded (US) thruster. Also, based on carbon deposition measurements, the erosion rates at the inner and outer walls of the MS thruster are found to be lower by at least 2300 and 1875 times, respectively. Erosion was so low along these walls that the rates were below the resolution of the profilometer. Using a sputtering yield model with an energy threshold of 25 V, the simulations predict a reduction of ~600 at the MS inner wall. At the outer wall ion energies are computed to be below 25 V, for which case we set the erosion to zero in the simulations. When a 50-V threshold is used the computed ion energies are below the threshold at both sides of the channel. Uncertainties, sensitivities and differences between theory and experiment are also discussed.

plasma simulations↗

NASA NTRS1972

Effects of the hollow cathode discharge on throttled ion thruster performance.

An experimental study performed on a 20 cm electron bombardment thruster shows substantial improvements in throttled performance can be achieved if cathode discharge region diameter and length, cathode flow rate and baffle aperture area are reduced simultaneously. This improvement is related to a thruster operating mode in which arc current does not vary significantly with arc voltage. In this mode it is possible to operate the thruster at a high arc voltage and a correspondingly high primary electron energy without the penalty of a prohibitively high arc power. Langmuir probe data indicate the arc current is limited by the cathode region plasma potential and electron density which are both nearly invariant with arc voltage.

Wilbur, P. J.↗

NASA NTRS1972

Magnetogasdynamic heat transfer to a hemispherical body in a supersonic argon plasma.

Measurements of the heat transfer to a hemispherical body in a Mach 4.6, low density, argon plasma are reported. The experiments demonstrated that local wall currents affected the heat flux to the electrically conducting surface, both without and with internally applied magnetic fields. Such currents were induced by a nonuniform shock-layer plasma potential; but with applied fields, Hall currents dominated. Magnetic fields significantly increased the stagnation-region heat flux while causing a reduction over the remainder of the hemisphere. For an electrically nonconducting surface, the local heat flux was reduced, and magnetic fields insignificantly affected the heat flux. Analytical treatments indicate that joule and radiation heating were negligible.

Nowak, R. J.↗

NASA NTRS1973

Interaction of the solar wind with the outer planets.

The hypersonic analog for the interaction of the solar wind with Jupiter, Saturn, Uranus, Neptune, and Pluto is used to provide estimates of shock shapes and locations, as well as average magnetosheath and/or ionosheath properties for these planets. Several representative spacecraft flyby trajectories (designed for outer-planet 'Grand Tour' simulations) are superimposed upon a series of figures in order to provide estimates of potential plasma and field parameters which may be encountered. Consideration is given first to the possibility that several of these planets have intrinsic magnetic fields and, secondly, to the interaction of the solar wind directly on the ionosphere should there actually be no intrinsic field. Saturn and Pluto are chosen as examples of this latter case.

Dryer, M.↗

NASA NTRS1976

The Skylab barium plasma injection experiments. II - Evidence for a double layer

Television observations of a barium-plasma flux tube extending from near 4500 km to near 10,000 km during a magnetic substorm and dawn-sector auroral display indicated several interesting anomalous events. Beyond 5500 km, there was a rapid increase in brightness accompanied by flux-tube splitting and diffusion, leaving behind a truncated single flux tube. From the orientation of the flux tube compared with theoretical field models, the presence of a substantial field-aligned current sheet is deduced. A suggested explanation of these phenomena is given in terms of a plasma potential double layer.

Wescott, E. M.↗

NASA NTRS1978

Measurement of non-Maxwellian electron velocity distributions in a reflex discharge

The results of a ruby laser Thomson scattering study of the space and time-resolved electron velocity distributions in a pulsed Penning discharge in hydrogen are presented. Electron densities were to the order of 10 to the 13th/cu cm and temperatures were roughly 3 eV. This point is just prior to the cessation of the discharge ohmic heating pulse. For magnetic strengths less than 200 G, Maxwellian distributions were found over an energy range six times thermal energy. Temperatures agreed with Langmuir probe data. For fields of 450 G, chaotic plasma potentials were observed to be unstable and the Thomson scattering showed that the electron velocity distributions had central temperatures of 2 eV and wing temperatures of 15-12 eV.

Phipps, C. R., Jr.↗

NASA NTRS1981

Additional extensions to the NASCAP computer code, volume 3

The ION computer code is designed to calculate charge exchange ion densities, electric potentials, plasma temperatures, and current densities external to a neutralized ion engine in R-Z geometry. The present version assumes the beam ion current and density to be known and specified, and the neutralizing electrons to originate from a hot-wire ring surrounding the beam orifice. The plasma is treated as being resistive, with an electron relaxation time comparable to the plasma frequency. Together with the thermal and electrical boundary conditions described below and other straightforward engine parameters, these assumptions suffice to determine the required quantities. The ION code, written in ASCII FORTRAN for UNIVAC 1100 series computers, is designed to be run interactively, although it can also be run in batch mode. The input is free-format, and the output is mainly graphical, using the machine-independent graphics developed for the NASCAP code. The executive routine calls the code's major subroutines in user-specified order, and the code allows great latitude for restart and parameter change.

Mandell, M. J.↗

NASA NTRS1982

A comparison of experimental and computer model results on the charge-exchange plasma flow from a 30 cm mercury ion thruster

Ion thrusters can be used in a variety of primary and auxiliary space-propulsion applications. A thruster produces a charge-exchange plasma which can interact with various systems on the spacecraft. The propagation of the charge-exchange plasma is crucial in determining the interaction of that plasma with the spacecraft. This paper compares experimental measurements with computer model predictions of the propagation of the charge-exchange plasma from a 30 cm mercury ion thruster. The plasma potentials, and ion densities, and directed energies are discussed. Good agreement is found in a region upstream of, and close to, the ion thruster optics. Outside of this region the agreement is reasonable in view of the modeling difficulties.

Gabriel, S. B.↗

NASA NTRS1987

Langmuir probe surveys of an arcjet exhaust

Electrostatic (Langmuir) probes of both spherical and cylindrical geometry have been used to obtain electron number density and temperature in the exhaust of a laboratory arcjet. The arcjet thruster operated on nitrogen and hydrogen mixtures to simulate fully decomposed hydrazine in a vacuum environment with background pressures less than 0.05 Pa. The exhaust appears to be only slightly ionized (less than 1 percent) with local plasma potentials near facility ground. The current-voltage characteristics of the probes indicate a Maxwellian temperature distribution. Plume data are presented as a function of arcjet operating conditions and also position in the exhaust.

Zana, Lynnette M.↗

NASA NTRS1987

Ground-based plasma contractor characterization

Presented are recent NASA Lewis Research Center (LeRC) plasma contractor experimental results, as well as a description of the plasma contractor test facility. The operation of a 24 cm diameter plasma source with hollow cathode was investigated in the lighted-mode regime of electron current collection from 0.1 to 7.0 A. These results are compared to those obtained with a 12 cm plasma source. Full two-dimensional plasma potential profiles were constructed from emissive probe traces of the contractor plume. The experimentally measured dimensions of the plume sheaths were then compared to those theoretically predicted using a model of a spherical double sheath. Results are consistent for currents up to approximately 1.0 A. For currents above 1.0 A, substantial deviations from theory occur. These deviations are due to sheath asphericity, and possibly volume ionization in the double-sheath region.

Patterson, Michael J.↗

NASA NTRS1987

Langmuir probe surveys of an arcjet exhaust

Zana, Lynnette M.↗

NASA NTRS1988

An Experimental Investigation of an Arcjet Thruster Exhaust Using Langmuir Probes

Electrostatic (Langmuir) probes of both spherical and cylindrical geometry have been used to obtain electron number density and temperature in the exhaust of a laboratory arcjet. The arcjet thruster operated on nitrogen and hydrogen mixtures to simulate fully decomposed hydrazine in a vacuum environment with background pressures less than 5 x 10 to the -2 Pa. The exhaust appears to be only slightly ionized (less than 1 percent) with local plasma potentials near facility ground. The current-voltage characteristics of the probes indicate a Maxwellian temperature distribution. Plume data are presented as a function of arcjet operating condition and also position in the exhaust.

Carney, Lynnette M.↗

NASA NTRS1988

Ground-based plasma contactor characterization

This paper presents recent NASA Lewis Research Center plasma contactor experimental results, as well as a description of the plasma contactor test facility. The operation of a 24-cm-diameter plasma source with hollow cathode was investigated in the 'ignited-mode' regime of electron current collection from 0.1 to 7.0 A. These results are compared to those obtained with a 12-cm plasma source. Full two-dimensional plasma potential profiles were constructed from emissive probe traces of the contactor plume. The experimentally measured dimensions of the plume sheaths were then compared to those theoretically predicted, using a model of a spherical double sheath. Results are consistent for currents up to approximately 1.0 A. For currents above 1.0 A, substantial deviations from theory occur. These deviations are due to sheath asphericity, and, possibly, volume ionization in the double-sheath region.

Patterson, Michael J.↗

NASA NTRS1989

The Space Station Photovoltaic Panels Plasma Interaction Test Program: Test plan and results

The Plasma Interaction Test performed on two space station solar array panels is addressed. This includes a discussion of the test requirements, test plan, experimental set-up, and test results. It was found that parasitic current collection was insignificant (0.3 percent of the solar array delivered power). The measured arcing threshold ranged from -210 to -457 V with respect to the plasma potential. Furthermore, the dynamic response of the panels showed the panel time constant to range between 1 and 5 microsec, and the panel capacitance to be between .01 and .02 microF.

Nahra, Henry K.↗

NASA NTRS1990

The Space Station photovoltaic panels plasma interaction test program - Test plan and results

The plasma Interaction Test performed on two space station solar array panels is addressed. This includes a discussion of the test requirements, test plan, experimental set-up, and test results. It was found that parasitic current collection was insignificant (0.3 percent of the solar array delivered power). The measured arcing threshold ranged from -210 to -457 V with respect to the plasma potential. Furthermore, the dynamic response of the panels showed the panel time constant to range between 1 and 5 microsec, and the panel capacitance to be between .01 and .02 microF.

Nahra, Henry K.↗

NASA NTRS1991

A preliminary characterization of applied-field MPD thruster plumes

Electric probes, quantitative imaging, and emission spectroscopy were used to study the plume characteristics of applied field magnetohydrodynamic thrusters. The measurements showed that the applied magnetic field plays the dominant role in establishing the plume structure, followed in importance by the cathode geometry and propellant. The anode radius had no measurable impact on the plume characteristics. For all cases studied the plume was highly ionized, though spectral lines of neutral species were always present. Centerline electron densities and temperatures ranged from 2 times 10 (exp 18) to 8 times 10 (exp 18) m(exp -3) and from 7500 to 20,000 K, respectively. The plume was strongly confined by the magnetic field, with radial density gradients increasing monotonically with applied field strength. Plasma potential measurements show a strong effect of the magnetic field on the electrical conductivity and indicate the presence of radial current conduction in the plume.

Myers, Roger M.↗