Myron J. Mandell

Elastic Scattering of Ions in Electrostatic Thruster Plumes

It is postulated that the ion population and energy in most ion and Hall-Effect thruster plumes, at angles beyond the main beam divergence, is largely determined by the elastic scattering of high-energy ions by neutral atoms. A theoretical model of the scattered ion density and energy is presented. The model accounts for the presence of a main ion beam emanating from the thruster and neutral atoms from the ambient and thruster-induced environment. The treatment of elastic scattering incorporates angle-dependent differential cross sections that are computed classically. Results from the model are compared with ion flux and energy measurements taken by a retarding potential analyzer as a function of angle from the thruster’s axis of symmetry. The good agreement between model results and measurements for ion energy and flux suggests that the determination of the differential cross section using classical formulations is adequate for the high-energy ions associated with electrostatic propulsion plumes. But the comparisons also emphasize the importance of accurate models of the neutral particle densities in these plumes. Nomenclature A = area of thruster acceleration channel, m 2 E = impact energy associated with the scattering event, eV F = ion particle flux, particles/m 2 s Fr = radial component of ion flux vector, particles/m 2 s Fz = axial component of ion flux vector, particles/m 2 s f = multiple ionization flux fraction I = differential cross section, m 2 /sr J = current, A M = reduced mass, kg m = particle mass, kg ˙ mi = ion mass flow rate, kg/s N = number of scattered particles n = ion particle density, particles/m 3 n0 = neutral particle density, particles/m 3 n∞ = reference ion particle density, particles/m 3

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.

Ion Engine Plume Interaction Calculations for Prototypical Prometheus 1

Prometheus 1 is a conceptual mission to demonstrate the use of atomic energy for distant space missions. The hypothetical spacecraft design considered in this paper calls for multiple ion thrusters, each with considerably higher beam energy and beam current than have previously flown in space. The engineering challenges posed by such powerful thrusters relate not only to the thrusters themselves, but also to designing the spacecraft to avoid potentially deleterious effects of the thruster plumes. Accommodation of these thrusters requires good prediction of the highest angle portions of the main beam, as well as knowledge of clastically scattered and charge exchange ions, predictions for grid erosion and contamination of surfaces by eroded grid material, and effects of the plasma plume on radio transmissions. Nonlinear interactions of multiple thrusters are also of concern. In this paper we describe two- and three-dimensional calculations for plume structure and effects of conceptual Prometheus 1 ion engines. Many of the techniques used have been validated by application to ground test data for the NSTAR and NEXT ion engines. Predictions for plume structure and possible sputtering and contamination effects will be presented.

Nascap-2k Spacecraft-Plasma Environment Interactions Modeling: New Capabilities and Verification

Nascap-2k is a three-dimensional computer code that models interactions between spacecraft and plasma environments in low-Earth, geosynchronous, auroral, and interplanetary orbits. The code builds on physical principles, mathematical algorithms, and user experience developed over three decades of spacecraft charging research. Nascap-2k has improved numeric techniques, a modern user interface, and a simple, interactive satellite surface definition module (Object ToolKit). 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. Previously, we reported on the accuracy of Nascap-2ks charging and current collection calculations by comparing computed currents and potentials with analytic results, and by comparing Nascap-2k results with published calculations using the earlier lower resolution codes, NASCAP/GEO, NASCAP/LEO, and POLAR. Here we examine the accuracy and limitations of two new capabilities of Nascap-2k: modeling of plasma plumes such as generated by electric thrusters and enhanced PIC computational capabilities. Nascap-2k models one or more ion engine plumes in full three-dimensional geometry, including plume-plume and plume-spacecraft interactions. The primary thruster beam, parameters describing the neutral efflux, and the initial charge-exchange plume are imported from a PlumeTool-generated file. Nascap-2k generates and tracks charge-exchange Ions to obtain plasma densities and calculates potentials consistent with plasma densities and object surfaces. Nascap-2ks PIC capability has been expanded to include boundary injection, particle splitting, and substep charge deposition. The boundary injection replaces collected particles in long running calculations. The particle splitting allows for modeling the effects of the thermal distribution of velocities, as well as accommodating particle weight to variable grid cell volume. The substep charge deposition makes possible calculations in which two effects have significantly different timescales. We use calculations for simple geometries to explore the accuracy and limitations of these capabilities.

Extrapolation of electrical breakdown currents from the laboratory to Space Station

Recent experiments conducted in a plasma chamber at NASA/MSFC on anodized aluminum coatings representative of Space Station Freedom design show that if the aluminum used as a thermal control coating is biased more than 80 V negative with respect to the plasma, the anodization will experience dielectric breakdown. As the thin anodization layer creates a capacitive charge buildup, large currents are observed during the arc. How plasma generation at the arc site can support large currents and discharge the surface charge layer is investigated. The importance for Space Station Freedom is that currents similar to those observed in the laboratory can be observed on orbit.

Simulation of Auroral Charging in the DMSP Environment

The Earth’s aurora provides one of the two regions where spacecraft charging is of greatest concern, along with the geosynchronous orbit. Nascap-2k is designed to model charging in both regions however the specific numerical settings for each are very different. This paper describes the physical models implemented in Nascap-2k and can serve as a guide to using the code for auroral charging simulations. Simple test objects are used to illustrate the numerical options. A full scale simulation of a satellite of interest is also included. The Air Force DMSP (Defense Meteorological Satellite Program) satellites have been flying in the auroral region for almost three decades and provide a wealth of environmental data as well as measurements of the vehicle charging.

The Electric Propulsion Interactions Code (EPIC): A Member of the NASA SEE Program Toolset

Science Applications International Corporation is currently developing the Electric Propulsion Interactions Code, EPIC, as part of a project sponsored by the Space Environments and Effects Program at NASA Marshall Space Flight Center. Now in its second year of development, EPIC is an interactive computer toolset that allows the construction of a 3-D spacecraft model and the assessment of a variety of interactions between its subsystems and the plume from an electric thruster. This paper reports on the progress of EPIC including the recently added ability to exchange results with the spacecraft charging analyzer program, Nascap2k. The capability greatly enhances EPIC’s range of applicability. Expansion of the toolset’s various physics models proceeds in parallel with the overall development of the software. Also presented are recent upgrades of the elastic scattering algorithm in the electric propulsion Plume Tool. These upgrades are motivated in part by the need to assess the impact of elastically scattered ions on spacecraft for ion beam energies that exceed 1000 eV. Such energy levels are expected in future high-power (>10 kW) ion propulsion systems empowered by nuclear sources.

Modeling of DMSP Surface Charging Events

Measured spectra from a selection of Defense Meteorological Satellite Program-observed charging events were analyzed, with the goal of improving modeling of auroral charging events. Each data set was examined for correlations between the chassis potential and the measured fluxes. Chassis potential was found to be correlated with both the median electron energy and the net electron flux. The background plasma density cannot be determined directly from the measured spectra due to distortions from charging and the variation in sensitivity with incident particle energy across each energy channel. Simulations using time-varying spectra show that the chassis potential responds to even fairly subtle environment changes on a millisecond time scale. As the environment measurements take 1 s per spectrum, the chassis potential can easily be changing, either steadily or erratically, during that time. Once the limitations of the measurements are accounted for, the calculated chassis potential agrees with the measured potential through the modeled charging events.

NASCAP-2K SIMULATIONS OF SPACECRAFT CHARGING IN TENUOUS PLASMA ENVIRONMENTS

Nascap-2K is being developed under sponsorship of the US Air Force Research Laboratory and NASA’s Space Environment Effects (SEE) program. It replaces the twenty-five year old NASCAP/GEO code, along with other aging spacecraft plasma interactions codes such as NASCAP/LEO and POLAR. Nascap-2K includes an interactive Object Toolkit (OTk) for spacecraft surface definition, uses a newly developed Boundary Element Method (BEM) treatment for surface charging, and uses advanced numeric techniques developed for the DynaPAC code for external potentials, charge densities, and particle trajectories. In this paper we give a brief description of the interface and some of the numerical techniques, and show simulation results for a geosynchronous satellite (DSCS-IIII), a satellite in the Solar Wind (STEREO), and a satellite in Mercury orbit (MESSENGER).