Modeling Of Spacecraft Charging Research Articles

Strategies for Determining Electron Yield Material Parameters for Spacecraft Charge Modeling

AbstractAccurate modeling of spacecraft charging is essential to mitigate well‐known and all‐too‐common deleterious and costly effects on spacecraft resulting from charging induced by interactions with the space plasma environment. This paper addresses how limited availability of electron emission and transport properties of spacecraft materials—in particular, secondary electron yields—and the wide ranges measured for such properties pose a critical issue for modeling spacecraft charging. It describes a materials charging database for electron emission properties under development, which facilitates more accurate spacecraft charge modeling when used in concert with the strategies outlined herein. These data and techniques provide tools for more accurate material selection, increased confidence in charge models, and a concomitant decrease in mission risk. They also allow better customization of models in response to prolonged space environment exposure and specific mission requirements, which may evolve materials properties.

Modeling of Spacecraft Charging Dynamics Using COULOMB-2 Code

This paper describes modeling of spacecraft charging dynamics which is used in COULOMB-2 code in the case of spacecraft surface complex shape. The modeling of spacecraft charging is carried out via numerically solving the system of differential equations for time variations of local electric charge on every discrete element of the spacecraft surface. The presented computation results are obtained for spacecraft charging in hot magnetosphere plasma for several spacecraft design elements in a time interval of 20–10 000 s. The results are compared with the similar ones obtained with the NASCAP-2k and MUSCAT codes, and a good consistency was found.

Perspectives on the Distributions of ESD Breakdowns for Spacecraft Charging Applications

Electrostatic discharge (ESD) continues to pose significant risks to space missions despite decades of intense study. Tabulated values of material breakdown strength used in spacecraft charging models are often based on cursory measurements that may not be fully relevant to a given mission. Materials physics offers insight into the pertinent variables that affect breakdown and how to address them experimentally for spacecraft applications. We present measured distributions of ESD data across several test configurations for three polymeric materials that, taken together, begin to provide an understanding of how to estimate the likelihood of ESD events over a spacecraft’s mission lifetime. We discuss how consequences of these results apply to spacecraft charging modeling and design considerations.

Comparison of Coulomb-2, NASCAP-2K, MUSCAT and SPIS codes for geosynchronous spacecraft charging

In the paper, we discuss numerical modeling of spacecraft charging in geosynchronous orbit which was done using Coulomb-2, NASCAP-2K and SPIS codes, and present data obtained using the MUSCAT code too. The main distinctive features of the codes above are specified, description of the Coulomb-2 code is given in more details. We compare the results of calculations obtained for similar models of spacecrafts with identical surface materials and parameters of the space plasma environment. The possible reasons of divergence of the modeling results are discussed.

Macroscopic model of internal discharging in polymer under electron beam irradiation

Internal charging effect and discharging effect in a dielectric material are one of the key factors threatening the spacecraft safety. Most of the spacecraft charging models could calculate only the voltage variation and electric field distribution in the spacecraft internal charging process, without estimating the consequence of discharging, such as the magnitude of voltage drop and discharging current. In this article, we present a combined model which could calculate both the charging effect and the discharging effect on polymer. The model is verified by the experimental results, and it shows that the residual voltage after discharging and the pulse width of discharging current grow with the increase of polymer resistivity. The magnitude of discharging current is determined by the breakdown electrical field.

New Frontiers in Spacecraft Charging

Spacecraft charging, as a field, is continually being recharged by new developments in understanding, new materials and technologies, and new approaches to both new and old problems. I will discuss some of the new frontiers in understanding spacecraft charging in this paper, as well as referencing relevant papers from the 11th Spacecraft Charging Technology Conference. Spacecraft charging is highly material-property dependent. For example, the secondary electron emission, photoemission, bulk electrical resistivity, and surface resistivity are important parameters that help determine the extent of spacecraft charging (both on the surface and inside the spacecraft) in any given environment. Our understanding of these material properties is one of the new frontiers in spacecraft charging. I will discuss how these fundamental material properties have been found to depend on the following: proper measurement techniques, temperature, radiation flux, electric field, surface treatment, surface contamination from plumes and outgassing, surface modification through arcing and vacuum exposure, and synergistic effects. Modeling of spacecraft charging is a second new frontier. New developments in modeling have both improved our understanding of spacecraft charging and enabled us to model situations that are dynamic and geometrically complex. New schemes for treating both space and time variations of fields, particle fluxes, and spectra have made our modeling more precise and accurate. Now, many more spacecraft are being launched into low Earth orbit and the radiation belts. Modeling charging effects more accurately in those orbits will become more important than ever. A third new frontier in spacecraft charging is novel mitigation techniques. Surface materials and simple passive devices that emit electrons as fast as they are collected seem to make real-time charge mitigation cheaply and reliably achievable for the first time. Novel solar cell configurations and coverglass materials promise to make arcing, both of the primary electrostatic discharge (ESD) type and sustained arcing between cells or strings, a thing of the past. Superconducting cables may obviate the high voltages that lead to arcing. New cooperation between spacecraft and solar array manufacturers and spacecraft charging experts may help to prevent the spacecraft charging mistakes of the past. Furthermore, the final frontier is dealing with new materials and higher power requirements. Lightweight spacecraft materials are, in some cases, prone to exacerbate charging or arcing and may allow transmission of electromagnetic interference into sensitive electronics. New solar cell active materials may increase the effects of arcing on solar cell and solar array performance, even for primary ESD events. Higher power requirements may require longer transmission cables, which may increase the need for higher voltages, making arcing more likely. If superconducting cables become a reality, magnetics may become very important for spacecraft control and stability. What will happen to a superconducting cable if it must carry the increased current in an arcing event of very short duration?

Validation of the Plasma Densities and Temperatures From the ISS Floating Potential Measurement Unit

The validation of the floating potential measurement unit (FPMU) plasma density and temperature measurements is an important step in the process of evaluating International Space Station (ISS) spacecraft charging issues including vehicle arcing and hazards to crew during extravehicular activities. The highest potentials observed on the Space Station are due to the combined Vsp times B effects on a large spacecraft and the collection of ionospheric electron and ion currents by the 160-V U.S. solar array modules. The ionospheric plasma environment is needed for input to the ISS spacecraft charging models used to predict the severity and frequency of occurrence of ISS charging hazards. The validation of these charging models requires the comparison of their predictions with measured FPMU values. The FPMU measurements themselves must also be validated for use in manned flight safety work. This paper presents preliminary results from a comparison of densities and temperatures derived from the FPMU Langmuir probes and plasma impedance probe with the independent density and temperature measurements from a spaceborne ultraviolet imager, a ground-based incoherent scatter radar, and ionosonde sites.

Charging of the International Space Station as Observed by the Floating Potential Measurement Unit: Initial Results

The floating potential measurement unit (FPMU) is a multiprobe package designed to measure the floating potential of the International Space Station (ISS) as well as the density and temperature of the local ionospheric plasma environment. The purpose of the FPMU is to provide direct measurements of ISS spacecraft charging as continuing construction leads to dramatic changes in ISS size and configuration. FPMU data are used for refinement and validation of the ISS spacecraft charging models used to evaluate the severity and frequency of occurrence of ISS charging hazards. The FPMU data and the models are also used to evaluate the effectiveness of proposed hazard controls. The FPMU consists of four probes: a floating potential probe, two Langmuir probes, and a plasma impedance probe. These probes measure the floating potential of the ISS, plasma density, and electron temperature. Redundant measurements using different probes support data validation by interprobe comparisons. The FPMU was installed by ISS crew members during an extra-vehicular activity on the starboard (S1) truss of the ISS in early August 2006 when the ISS configuration included only one 160-V U.S. photovoltaic (PV) array module. The first data campaign began a few hours after installation and continued for over five days. Additional data campaigns were completed in 2007 after a second 160-V U.S. PV array module was added to the ISS. This paper discusses the general operational characteristics of the FPMU as integrated on ISS, the functional performance of each probe, the charging behavior of the ISS before and after the addition of a second 160-V U.S. PV array module, and initial results from model comparisons.

Forecasting electron precipitation based on predicted geomagnetic activity

We have developed a statistical model of precipitating electrons based on 10 years worth of NOAA-12 Space Environment Monitor (SEM) data. For each hour of UT, the electron spectral shape is defined for solar activity (using F 10.7), geomagnetic activity (using Dst, Kp, and PC), invariant latitude (40 to 90°N and −40° to −90°S), and every hour of magnetic local time to form a climatology. The current version of our climatology includes a description of the precipitating differential number flux, unfolded from the integral high-energy telescopes of SEM and its lower-energy differential measurements. In total, the energy range covered by SEM is from 300 eV to 1 MeV. We use real-time Dst predictions from several sites on the Internet which forecast Dst up to 1 h ahead of time to predict the corresponding precipitating electron flux. Our electron precipitation forecast contributes to space weather applications by predicting a spectral description of atmospheric electron energy input. These electron energy spectra can be used to generate atmospheric energy input, ionization rates and height-integrated Hall and Pedersen conductivities, which are important parameters to models such as AMIE and TIME-GCM. Our data could also improve the space environment description used in spacecraft charging models.