Adrian Wheelock

Nascap-2k Spacecraft Charging Code Overview

Nascap-2k is a modern spacecraft charging code, replacing the older codes NASA Charging Analyzer Program for GEosynchronous Orbit (NASCAP/GEO), NASA Charging Analyzer Program for Low-Earth Orbit (NASCAP/LEO), Potentials Of Large objects in the Auroral Region (POLAR), and Dynamic Plasma Analysis Code (DynaPAC). The code builds on the physical principles, mathematical algorithms, and user experience developed over three decades of spacecraft charging research. Capabilities include surface charging in geosynchronous and interplanetary orbits, sheath, and wake structure, and current collection in low-Earth orbits, and auroral charging. External potential structure and particle trajectories are computed using a finite element method on a nested grid structure and may be visualized within the Nascap-2k interface. Space charge can be treated either analytically, self-consistently with particle trajectories, or consistent with imported plume densities. Particle-in-cell (PIC) capabilities are available to study dynamic plasma effects. Auxiliary programs to Nascap-2k include Object Toolkit (for developing spacecraft surface models) and GridTool (for constructing nested grid structures around spacecraft models). The capabilities of the code are illustrated by way of four examples: charging of a geostationary satellite, self-consistent potentials for a negative probe in a low-Earth orbit spacecraft wake, potentials associated with thruster plumes, and PIC calculations of plasma effects on a very low frequency (about 1 to 20 kHz) antenna

Pulsed Plasma Thruster Plume Investigation Using a Current-Mode Quadruple Probe Method

A current-mode quadruple-Langmuir-probe method was developed and used to measure electron temperature, electron density, and the ratio of ion speed to most probable thermal speed in the plume of a rectangular-plate laboratory model Teflon ® pulsed plasma thruster. The current mode involves biasing all the electrodes and measuring the collected currents. The current collection theory and uncertainty analysis are presented. The thruster was operating at discharge energies of 5, 20, and 40 J and 15-microsecond pulses. Quadruple-probe measurements were taken at 10, 15, and 20 cm downstream of the Teflon propellant surface and angular locations of up to 40 deg off the centerline on planes perpendicular and parallel to the thruster electrodes. The electron temperature at 10 cm reaches values that range from 10 to 18 eV during the rise of the discharge current but remains below 5 eV for the rest of the pulse. The electron temperature shows no angular variation but reduces with increasing radial distance. The maximum electron density at 10 cm is 10, 20 7 × 10, 20 and 1.5 × 10 21 m −3 for the 5, 20, and 40-J discharge energy, respectively. The electron density reduces with increasing angle especially at larger downstream distances. The ion-speed ratios indicate supersonic ions and increase with downstream distance and for all angles considered. Ion speeds at 10 cm are approximately 30,000 m/s at the beginning of the pulse and reduce in magnitude to below 10,000 m/s near the end of the pulse.

Comparison of Low Earth Orbit Wake Current Collection Simulations Using Nascap-2k, SPIS, and MUSCAT Computer Codes

The structure of the wake generated by an object immersed in a dense, low temperature, drifting plasma is simulated using three different spacecraft charging software tools, Nascap-2k, the Spacecraft Plasma Interaction Software (SPIS), and the Multi-Utility Spacecraft Charging Analysis Tool (MUSCAT). Each tool uses different algorithms to simulate the particle dynamics, the space charge effect on the electric field, and the currents collected by the object. The system modeled is a plate with a high negative potential on the wake side. The results from the different simulations agree with each other and with experiments conducted on the same configuration. In particular, the nontrivial shape of the collected current density map is correctly simulated by all the codes. The comparison illustrates the strengths of the various approaches.

Charge Control of Geosynchronous Spacecraft Using Field Effect Emitters

Abstract : *Space Systems Loral is conducting an lR&D program to determine the feasibility and effectiveness of field effect electron emitters for potential control of geosynchronous altitude spacecraft. This electron emitters will be based on Spindt Cathode Field Emission Array Technologies. The configuration studied here consists of two emitters, each with an area of about 1 cm2 and emitting up to 1 mA of electrons at approximately 50eV energy. We show that it appears feasible to use electron emitters to control the surface charge of a satellite. Results concerning the placement and effectiveness of emitters and the spacecraft potential configuration under substorm conditions with and without emitter operations in sunlight, in eclipse, and during eclipse exit.

Investigation of ion beam neutralization processes with 2D and 3D PIC simulations

Abstract While it is common knowledge that ion beams are easily neutralized for both current and charge density using a variety of means, the precise process of neutralization remains unknown. With the increasing importance of electric propulsion, and in particular micropropulsion systems, this question is of significant importance. Additionally, it has bearing on thruster design, space instrument calibration, electrodynamic tethers, and ionospheric research. A review of the present state of knowledge on this topic is presented as well as results from ion beam simulations using 2D and 3D Particle-in-Cell (PIC) codes. We investigate both the early “filling” problem of the beam starting to move away from the spacecraft and the steady state problem where the beam encounters a wall at an infinite distance from the spacecraft.

Ion Beam Neutralization Processes for Electric Micropropulsion Applications

Abstract : While it is common knowledge that ion beams are easily neutralized using a variety of means, the precise process of neutralization remains unknown. With the increasing popularity of electric propulsion, and in particular micropropulsion systems, this question is of significant importance. Additionally, it has a bearing on thruster design, space instrument calibration, electrodynamic tethers, and ionospheric research. A review of the present state of knowledge on this topic is presented as well as results from ion beam simulations using 2D and 3D Particle-in-Cell codes. The grid generation methodology, adaptation, charged-particle transport, and field solver methodologies of the 3D code are reviewed. The simulations show electrons moving to neutralize the ion beam from background and neutralizer sources. The simulations show the dependence of neutralization on beam energy and the electron/ion velocity ratio. The results are compared favorably with previous computations and experimental observations.

AFRL Round-Robin Test Results on Plasma Propagation Velocity

The speed plasma propagates across a charged solar panel after a primary arc is one of the most important, yet poorly known, quantities in determining Electrostatic Discharge (ESD) currents for spacecraft arcing events. A review of the literature over the last two decades reveals that measured propagation velocity varies by as much as an order of magnitude. To overcome this deficiency, a round-robin set of tests was initiated with partners from industry, academia, NASA and the U.S. Air Force. This paper will provide the most recent results from the Air Force Research Laboratory testing conducted at the Spacecraft Charging and Instrument Calibration Laboratory.

Preliminary Measurements of ESD Propagation on a Round Robin Coupon

The propagation dynamics of spacecraft electrostatic discharge (ESD) flashover plasmas have been a topic of increasing interest in the past few years. To investigate ESD propagation and possible contributions from facilities, methodology, and analysis, we performed inverted gradient ESD tests of an International Space Station solar array coupon as part of a U.S.-wide Round-Robin test campaign. Geosynchronous earth orbit charging environments were simulated and current transients were simultaneously measured on ten solar array strings. In addition, images of ESD events were collected by low- and high-speed cameras, and an array of eight Langmuir probes positioned above the coupon was monitored. Preliminary results will be discussed, with special attention given to the effect of different analysis techniques on derived propagation velocity. Comparisons will be made with the results from other labs, and implications for spacecraft ESD events will be discussed.

ELECTRON-ION BEAM COUPLING THROUGH COLLECTIVE INTERACTIONS

Abstract : Neutralization of ion beams in electric propulsion applications is a well-known phenomenon. The physics behind the robust matching of both ion and electron currents and densities are not. With electric propulsion devices moving into micro and macro regimes with colloids, FEEPs, and thruster arrays, thruster-neutralizer interactions are under increasing scrutiny. It is shown that Coulomb collisions, which can act to match velocities through strong ion-electron collisions between particles with low relative velocities, are far too slow to explain the phenomenon. Further examination of the strong beam-plasma, or Buneman, instability yields a candidate for the neutralization mechanism. Differences in simulations from analytic theory are discussed.

The Compact Environmental Anomaly Sensor Risk Reduction: A Pathfinder for Operational Energetic Charged Particle Sensors

Compact environmental anomaly sensor risk reduction (CEASE-RR) is a new sensor designed for anomaly attribution due to the space radiation environment. It does this using two solid-state particle telescopes that have been designed to measure proton and electron fluxes that are the drivers for three of the four primary space environment effects (event total dose, deep-dielectric charging, and single event effects). These telescopes are integrated into a compact package along with space reserved for a planned electrostatic analyzer being developed for the final CEASE 3 design (covering the fourth primary space environment effect—surface charging). The sensors themselves will measure a wider dynamic range in particle flux, provide higher energy resolution, have better out-of-band contamination rejection, and improved diagnostic capability compared to previous CEASE instruments. The CEASE-RR instrument is planned to be launched in 2018 to geostationary orbit as part of an Air Force Research Laboratory flight experiment. The sensor design, calibration, and planned flight experiment objectives are described in this paper.