Abstract

A three-year program to develop a Direct Drive Hall Effect Thruster (D2HET) system has begun as part of the NASA Advanced Cross-Enterprise Technology Development initiative. The system is expected to reduce significantly the power processing, complexity, weight, and cost over conventional low-voltage systems. The D2HET will employ solar arrays that operate at voltages greater than 300V, and will be an enabling technology for affordable planetary exploration. It will also be a stepping-stone in the production of the next generation of power systems for Earth orbiting satellites. This paper reports on initial D2HET high voltage solar array concepts, numerical models of their interactions with HET thruster plumes, and plans for plasma chamber verification tests. INTRODUCTION The empowerment of Hall thruster systems directly from advanced solar arrays offers considerable benefits over conventional systems. By eliminating, or significantly reducing, the power processing and distribution complexity, weight, and costs over conventional, low-voltage systems can be lowered to attractive levels for a variety of NASA's deep space missions. Such technologies may also have a wide range of applications to the commercial satellite industry. A major concern in these systems is the potentially destructive consequences of high-voltage arcs. Nearly every satellite manufacturer has experienced design problems due to high-voltage discharges as they scale up to higher power and voltages. 1 Discharge damage to solar arrays on commercial GEO communications satellites for example, has cost more than a hundred million of dollars in losses and redesign expenses in recent years. NASA programs have also been affected. In contrast to the system considered here the Terra solar array operates at 127V; D2HET requires 300V. The D2HET program will go beyond fixes in current systems to an entirely new design. The experience that is required to make this leap in technology draws from successful programs conducted in the past that concentrated on the elimination of array failures on satellites. DEVELOPMENT METHODLOGY The current goal is to design a Direct Drive Hall Effect Thruster system that provides an overall cost savings of at least 30% over conventional HET systems, 60% reduction of PPU mass, and 90% cutback of the PPU heat rejection (and associated radiator area) requirements. The major technological advance will be a 300V (or higher) solar array that will operate efficiently in both HET generated and ambient space plasmas. Design considerations include, interconnect shielding from the plasma, array string layout patterns, spacing and grouting, isolation diodes, substrate structural makeup, and grounding. A comparison of a D2HET and conventional HET system is shown in Successful operation of arrays in vacuum at >300V does not necessarily guarantee operation in the space ambient and HET-generated conditions. In order to assess the effects on the arrays under different environments and mission scenarios, two satellite configurations and two HET designs are investigated using 2-D and 3-D simulation tools. The satellites are the Russian EXPRESS, a geosynchronus communications spacecraft, and the U.S. Deep Space-1. The SPT-100 and BPT-4000 thrusters have been chosen as the two representative HETs. The approach is to determine the HET environments that interact with the arrays and use the information to design and test arrays under these conditions. Express and Deep Space-1 The two satellite models were generated using the 3-D, spacecraft interactions code EWB (Environment Work Bench). In the EXPRESS model the solar array panels were arranged according to the designs in the Express-A reports, and were configured to conduct suntracking rotation for mission studies. SPT-100 and BPT-4000 Hall Effect Thrusters In the past, numerous efforts to simulate plumes from electrostatic thrusters have allowed for the development of a complete modeling package for accessing plume interactions with spacecraft subsystems. ENVIRONMENT DEFINITION Thruster Plume and Interactions with Model Spacecraft In order to determine the HET-induced environment near the arrays a combination of simulation tools was used: a 1-D model of the HET discharge chamber, a 2-D plume code, and the 3-D interactions code. The plume model is discussed in detail in reference 11. Briefly, the model consists of a Lagrangian primary beam algorithm and a Particle-in-Cell (PIC) solver for computing ion production from charge-exchange events. The primary beam is assumed to be a collisionless, singly-ionized, quasi-neutral plasma, expanding in a density-gradient electric field. The neutral gas density has two components in space: unionized beam particles and un-ionized neutralizer gas (from the hollow cathode). The beam of neutrals from the thruster is computed using an annular anode gas flow model with isotropic emission from the ring. The hollow cathode neutrals are assumed to have an isotropic emission at a constant temperature. For simulation of laboratory plumes, a third contribution to the neutrals is added based on the chamber background density, and is assumed to be uniform. Charge-exchange (CEX) is computed using a twodimensional, R,Z-geometry PIC code. The main beam ion densities computed by the Lagrangian calculations, and the neutral gas profile, are used as input during the calculation. The code solves Poisson's equation on a finite element grid and iterates until steady state CEX densities and density-gradient potentials are selfconsistent. The required conditions at the exit of the thruster are provided by a 1-D model of the acceleration channel in an SPT100 that produced good qualitative agreement with measured thruster performance in the past. 12 Two different simulation cases of plume maps were initially performed for the SPT100 onboard the Express A2 satellite, distinguished by two mass flow rate values: 4.9 and 5.3 mg/s for the thruster and 0.49 mg/s and 0.371mg/s for the cathode-neutralizer. The higher mass flow rate case is presented here. The computed plume maps (and trajectories) are illustrated in The standard 2-D plume model described above was also used in the past to explain the observed trends in the BPT4000 Hall-Effect thruster. [11] The 2-D maps of the thruster exhaust were "fed" into EWB. EWB's simulation capabilities include a variety of models for assessing interactions effects from electric propulsion plumes. The thruster ion flux at any point, i, on a surface j, due to plume component k, is calculated as follows: In equation The sputtering of a spacecraft surface j at point i, due to energetic ion impact from the thruster is calculated based upon the material sputtering rate, where, ijk Y is the sputter yield of the material on surface j. Depending upon the duration of thruster operation, t, the total surface erosion by direct plume impact, as well as re-deposition of sputtered particles onto other surfaces, is determined by computing a net erosion/deposition rate. The net rate is calculated as follows: for each spacecraft structure j the sputtering rates at all points are averaged to produce a source term S j R , at the centroid of that surface. This source term is then used to calculate a deposition rate at each of the grid points for all surfaces: where, ij Ω is the solid angle subtended by surface j at point i, and Θ is the angle between the normal of the depositing surface and a ray from the sputtered surface centroid to point i. The net rate, i R is then computed at each point by, If i R > 0 it is a deposition rate, if i R < 0, it is an erosion rate. This rate is then integrated over the mission duration to get a total number of deposited particles per square meter. The integrated value is determined by calculating an average rate and then multiplied by the mission duration. This accounts for time-dependent changes in spacecraft geometry (such as solar array rotation). The model for determining the induced torque Γ on the spacecraft during thruster operation accounts for contributions from the thrust and from exhaust impingement on surfaces: where, 0 r r r − = ∆ is the position vector of surface j or thruster T form a reference point with position vector, 0 r . The force j f is the momentum imparted to a surface from plume particles, per unit time. A choice to use either specular elastic reflection from the surface (colliding particle is reflected with the same speed and incidence angle equals the reflection angle), or diffuse reflection (based on material-dependent accommodation coefficients), is available. Comparisons with Measurements The results from the simulations are compared with measurements taken onboard the Express-A #2 and Express-A #3 satellites, during operation of the SPT100s. Measurements from two ion-flux sensors (DRT) onboard the Russian Express-A #2, were provided by the NASA Glenn Research Center. Data from one ion-flux sensor has been used, located at a distance of 1.352m from the thruster RT4, at an angle of approximately 80deg relative to the Xc axis. The second sensor is positioned under the MLI (multi-layer insulation) and did not provide any useful information. The sensor positions and thruster unit #4 (TU4) on the Express-A spacecraft are circled in As shown in Calculations similar to those shown above were also performed, for both Express and DS-1, using the BPT4000 plume map. D2HET Test Facility A vacuum chamber at MSFC is being configured for plasma testing of design concepts. This facility is shown in With the source on, the background pressure is in the high 10 -5 to low 10 -4 torr. Various diagnostics will be used to determine the plasma conditions around the array segments. Current collection by the solar array segment as a function of voltage (relative to the local plasma) will be measured and possible arcing will be detected. The surface voltage and sheath structure around solar cells will also be measured. The techniques to obtain these measurements are as follows: • Langmuir probe to measure the plasma density, electron temperature and local plasma potential in the vicinity of the solar array segment • Current/Voltage instrumentation oscilloscopes in conjunction with picoammeters or current amplifiers to measure the small currents collected by biased solar array segments and the applied voltage • Oscilloscope to trigger on current surges to the solar array to detect arcing on the segment • Video camera to view the solar array segment in the chamber during testing so that arcs on the segment can be captured by the video • Emissive probe will be swept in front of solar array segments during test to map the sheath structure extending into the plasma as a function of voltage on the segment, the array geometry and the plasma conditions • TREK probe-commercially available probe biases to zero electric field in order to provide non-contact surface voltage measurements. PRELIMINARY DESIGNS A preliminary set of candidate array designs has been selected for testing. During the initial tests mockups may be used for some designs. Several of the solar cell arrays will be based on the Space Station array design shown in ISS specs with 8cmX8cmX8mil cells This test article will either be a mock-up of the ISS solar array design or will be a segment supplied by Glenn Research Center. Preferably, a 4x5 cell array will be utilized to minimize the cell edge length around the edge relative to the length of cell adjacent to cell. ISS Specification with Solid or Mesh Cover of Cell Gap This will be the same configuration as in 2.2.1 except that the gap between cells will be covered by a grid. The electron collection is due to the fields from within the gap due to the high voltage solar cells. A grid that covers this gap will restrict the field and limit current collection from the local plasma. ISS Specification with RTV Grouting-2X2 and 4X4 Cell Segments This set will match specs of the ISS but with an RTV grouting applied to all gaps between the cells. This will reduce parasitic current collections and increase arc thresholds. Large Multi-cell Cover Glass An electrostatically clean solar array has been developed for scientific requirements. It utilizes a single large cover glass that covers a large number of solar cells at a time. This has the potential to significantly reduce high voltage cell exposure to the plasma and reduce interactions. This will be mocked up based on Composite Optics design shown in Large Cover Glass Overhang and Overlap ISS cell size will be utilized. A mocked up segment will allow examination of the effect of having a larger overhang of the cover glass. The ISS design is 3 mils. By increasing this to 10 or 20 mils and overlapping the cell edges by increasing alternating cells in height by 10 mils, the cell edges will overlap and prevent the electric field penetration into the plasma. Thin Film Arrays Glenn Research Center is developing several flexible thin film array technologies. An example is shown in Concentrators TECSTAR's concentrator shown below in SUMMARY This paper reported on the initial efforts of a NASA/GRC program to develop a Direct Drive Hall Effect Thruster system design. The goal is to advance power system technology to operate free of anomalies above 300V in a HET environment. The approach take is to; 1) Build on previous experience in D2HET demonstration programs at GRC; 2) Extend the knowledge gained during recent array anomaly analysis and mitigation efforts; 3) Define HET-satellite system configurations to baseline designs; 4) Use validated modeling tools to predict the plasma environments in which the high voltage components must operate freely of arcing and with minimal parasitic currents; 5) Stand-up a plasma test facility to test high voltage array concepts; 6) Use modeling to scale laboratory tests to high plasma density and temperate environments not attainable in the laboratory. The initial testing of mock-up and real array coupons is ongoing as of September 2001. We plan to report on initial interaction tests in upcoming conferences. ACKNOWLEDGMENTS