Gary Jongeward

A Hall effect thruster plume model including large-angle elastic scattering

A new model of the plasma plume from Hall Effect Thrusters (HETs) is presented. The model includes the self-expansion of the main beam by density gradient electric fields, lowenergy ions produced by resonant charge exchange between beam ions and neutral atoms (ambient and thruster-induced), and angle-dependent elastic scattering of beam ions off neutral atoms. The variation of radial velocities across the annular thruster beam is also included. The model is an advance over previous plume models in the way it numerically models the self-expansion of the main beam, and in particular, the treatment of elastic scattering using recently calculated differential cross sections. The results are compared with recent measurements of the energy and angledependent plume from the BPT4000 Hall-Effect Thruster. Both the intensity and energy dependence of the scattering peaks are compared. The principal result is that elastic scattering is the source of the majority of ions with energy greater than E/q=50V that are observed at angles greater than 45° with respect to the thrust axis. The model underscores the need for elastic scattering cross sections for multiply charged ions, as well as a better understanding of HET propellant utilization.

Electrical Breakdown Currents on Large Spacecraft in Low Earth Orbit

An experimental and theoretical investigation of an expanding plasma generated by an arc produced by biasing a conductor underneath a thin layer of anodized aluminum 160-V negative of a laboratory plasma that can produce large peak arc currents by discharging large surface areas is presented. A simple theory shows that the time scales and observed current magnitudes are consistent with the expansion of a discharge-generated plasma. The implication for large spacecraft in low Earth orbit, such as Space Station Freedom (SSF) which can store large amounts of charge, is that arcs with the same amount of energy similar to those observed in the laboratory may occur. The energy in these arcs degrade the surface of the anodized aluminum thermal control coatings by producing large pits in the surface. These pits tend to increase the temperature of the spacecraft, and the material from the pits can become an additional source of contamination . The rise time and intensity of theses arc could produce significant EMI. To prevent the occurrence of these undesirable effects, SSF will utilize a plasma contactor that will control the structure to ambient plasma potentials.

Threshold-determining mechanisms for discharges in high-voltage solar arrays

A theory is developed to account for the observed properties of discharges of solar arrays immersed in plasma. The theory is based on the assumption that a thin layer of insulating contaminant covers any metallic surface exposed to the plasma. Ions of the space plasma, attracted by the negative potential on the array, neutralize on the layers surface, resulting in a buildup of electric field in the layer. With continued charging, the internal field becomes sufficiently large to cause emission across the metal/layer interface and subsequent ionization and electron heating within the layer. The hot electrons can be emitted into vacuum; this emission constitutes the positive feedback mechanism leading to discharge. A mathematical description of the processes based on the foregoing assumptions is developed. Quantities derivable from the theory include a voltage threshold for arcing. The existence of the threshold and its predicted weak dependence on plasma density appear consistent with experimental results.

Plume Modeling of Stationary Plasma Thrusters and Interactions with the Express-A Spacecraft

Hall-effect thruster flight measurements are compared with results from two-dimensional plume and three-dimensional spacecraft interactions computer simulations. The measurements were acquired onboard Express-A 2 and A 3,two Russian communications satellites in geosynchronus orbit. The spacecraft carry four propulsion units for east-west and north-south station keeping. Each unit consists of two stationary plasma thrusters. Ion flux and energy spectra were recorded at various positions with respect to the thrusters and are compared with results from simulations using a uniform electron temperature, two-dimensional plume code that computes the expansion of the main ion beam by a fluid approach. The dynamics of the charge-exchange plasma are determined by a particle-in-cell method. Comparisons suggest good agreement for plume angles less than 40 deg and electron temperature between 8 and 11 eV. At approximately 4 and 9 m away from the thruster, and at plume angles less than 10 deg, the discrepancy between measured and computed values is found to be less than 10%. At larger angles, ion flux measurements exhibit large variations during operation of the same thruster. At 80 deg and 1.35 m away from the thruster, flux sensors recorded current densities that ranged between 12 and 55 mA/m 2 . The two-dimensional code computes 27 mA/m 2 for an anode mass flow rate of 5.3 mg/s at this location. Moments induced on the spacecraft during the operation of each thruster were also recorded by the attitude control system and are compared with results from a three-dimensional spacecraft interactions code. These measurements were taken during rotation of the solar arrays.

DIRECT DRIVE HALL THRUSTER SYSTEM DEVELOPMENT

The status of development of a Direct Drive Hall Thruster System is presented. In the first part, a study of the impacts to spacecraft systems and mass benefits of a direct-drive architecture is reviewed. The study initially examines four cases of SPT-100 and BPT-4000 Hall thrusters used for north-south station keeping on an EXPRESS-like geosynchronous spacecraft and for primary propulsion for a Deep Space-1 based science spacecraft. The study has also been extended to include the impact of direct drive on orbit raising for higher power geosynchronous spacecraft and on other deep space missions as a function of power and velocity change. The major system considerations for accommodating a direct drive Hall thruster are discussed, including array regulation, system grounding, distribution of power to the spacecraft bus, and interactions between current-voltage characteristics for the arrays and thrusters. The mass benefit analysis shows that, for the initial cases, up to 42 kg of mass savings is attributable directly to changes in the propulsion hardware. When projected mass impacts of operating the arrays and the electric power system at 300V are included, up to 63 kg is saved for the four initial cases. Adoption of high voltage lithium ion battery technology is projected to further improve these savings. Orbit raising of higher powered geosynchronous spacecraft is the mission for which direct drive provides the most benefit, allowing higher efficiency electric orbit raising to be accomplished in a limited period of time, as well as nearly eliminating significant power processing heat rejection mass. The total increase in useful payload to orbit ranges up to 278 kg (11%) for a 25 kW spacecraft, launched from an Atlas IIA. For deep space missions, direct drive is found to be most applicable to higher power missions with a velocity change up to several km/s, typical of several Discovery-class missions. In the second part, the status of development of direct drive propulsion power electronics is presented. The core of this hardware is the heater-keeper-magnet supply being qualified for the BPT- 4000 by Aerojet. A breadboard propulsion power unit is in fabrication and is scheduled for delivery late in 2003.

Model of Plasma Contactor Performance

A hollow cathode-based plasma contactor will bee own on the international spacestation to control the station’ s potential to within 40 V of the local ionosphere. Extensive testing of the plasma contactor has been conducted in vacuum facilitiesat theNASA LewisResearch Center. Signie cant performance differenceswereobserved between testsofthesameplasmacontactorindifferentfacilities.Whymeasuredplasmacontactorperformancediffersinthe laboratory in different tank environmentsand how the plasma contactor performance measured in the laboratory relatestoexpected performanceinspaceisaddressed. Presented aremodelsof plasmacontactorplasma generation and interaction in a laboratory environment, including anode area limiting. These models were integrated using the Space Station Environment Work Bench to predict plasma contactor operation, and the results are compared with the laboratory measurements. Nomenclature F = gas e ow rate, standard cubic centimeter per minute ID = total orie ce electron current, A Iemission = orie ce current emitted, A Ikeeper = keeper electrode current, A Iloss = ion loss rate, A Imax = maximum possible electron current, A Iprod = total ion production rate, A

Solar Arrays for Direct-Drive Electric Propulsion: Arcing at High Voltages

The results from an experimental investigation to assess arcing during operation of high voltage solar arrays in a plasma environment are presented. The experiments were part of an effort to develop systems that would allow safe operation of Hall-Effect Thrustefls) in direct-drive mode. Arc discharges are generated when the array is biased negative with respect to the plasma. If sustained for long periods of time between adjacent solar cells, arcs may severely damage a solar array, thus significantly shortening its lifetime. Most often sustained arcs are triggered by plasma produced during short-duration discharge arcs (approximately 20 microseconds). These trigger arcs are sparked between the semiconducting cell and the covering dielectric. Both trigger and sustained (greater than 1 millisecond) arcs have been captured during the tests. Current and voltage waveforms associated with the different arc events are presented. The test results have defined operational limits (thresholds) for the various array concepts studied that minimize the likelihood of damage from sustained arcs. Experimental trends regarding the effect of the solar array substrate on arc duration are also presented.

SPEAR 3 flight analysis: Grounding by neutral gas release, and magnetic field effects on current distribution

The Space Power Experiment Aboard Rockets (SPEAR) 3 experiment was launched on March 15, 1993, to test grounding devices for negative payloads. In this paper we review two aspects of the high-altitude flight data and compare them with preflight predictions. The SPEAR 3 neutral gas release experiment studied a grounding mechanism observed on previous flights during attitude control system (ACS) firings. Preflight calculations using Paschen law physics generalized to three dimensions predicted that the high rate gas release (about one order of magnitude below normal ACS) would reduce the rocket potential to within 200–300 V of plasma ground. The flight data is well fit by a value of −225 V. Orientation relative to Earths magnetic field had no effect on the floating potential or grounding operations but had a large effect on the portion of the current collected by the boom. We compare these flight measurements with preflight calculations made with the DynaPAC computer code.

Computer simulation of plasma electron collection by PIX-II

A wake model was defined for the NASCAP/LEO finite element model for the plasma interaction experiment (PIX-II) launched to study the interaction between high-voltage large solar arrays with the space plasma environment. The cell surface model considers the individual cells, distances between interconnects, and the fraction of surface covered by interconnects. Account is taken of the electrostatic potential around the spacecraft, which travels at 7500 mps, over five times the speed of thermal ions. Ram ions are produced ahead of the array and the wake ion density is described with a geometric shadowing model. The model correctly predicted the currents in high and low bias voltages when compared to orbital data. The panel snapover, however, was projected to occur at 100 V and instead occurred at 300 V, which indicates that the snapover state is bistable. Finally, a low potential was both predicted and measured in the wake.

Solar Arrays for Direct-Drive Electric Propulsion: Electron Collection at High Voltages

Solar array technologies that can empower electric thrusters in direct drive mode may provide significant mission benefits by reducing power processing, system complexity, weight, and cost over conventional systems. In direct-drive systems the solar arrays will operate at high voltages and must perform safely in the surrounding plasma environment, with minimal loss of performance due to parasitic current collection. For example, current state-of-the-art Hail effect thrusters in the kilowatt class require applied voltages of 300 V or higher. Results from experiments and modeling of electron current collection at high bias voltages (300-500 V) are presented. The experiments employed two sample solar array coupon technologies. A hollow cathode was used to emulate the induced environment around the solar arrays far from the Hall thruster and in nearby regions populated by charge-exchange plasma (10 1 2 -10 1 3 m - 3 and 0.5-1 eV). The measurements show that tens to hundreds of seconds are required before the collected current relaxes to a quasi-steady value. Comparisons with results from numerical calculations suggest that changes of the secondary electron yield properties of the dielectric materials may account for the observed current collection trends.