Spacecraft Plasma Interaction Research Articles

Severe Geostationary Environments: Numerical Estimation of Spacecraft Surface Charging from Flight Data

This paper defines a series of severe plasma environments in terms of spacecraft surface charging in geostationary orbit. In-flight plasma measurements are used to extract environments associated with large absolute spacecraft potentials and fluxes. An exhaustive list of events is selected over 15 years of Los Alamos National Laboratory data between the 1990s and 2000s. Electron spectra are compared to the literature and guidelines, especially the strong event measured by the Spacecraft Charging at High Altitude satellite in 1979. The impact on a telecommunications spacecraft of such charging environments is simulated numerically with the Spacecraft Plasma Interaction Software. The correlation between the model results and the instrument data indicates that the numerical approach is reliable for estimating charging risks in operational conditions. Measured environments are ranked on their susceptibility to generate significant charging and electrostatic discharge risks.

SPIS 5.1: An Innovative Approach for Spacecraft Plasma Modeling

Version 5.1 of Spacecraft Plasma Interaction Software (SPIS) has been recently released and is freely available for download to the whole community. It is the outcome of European Space Agency studies, SPIS-Geostationary Earth Orbit (GEO) and SPIS-SCI, and several contributions from the Spacecraft Plasma Interaction Network in Europe community members. It also relies on a Centre national d’etudes spatiales-sponsored activity to identify user needs of the industry for charging analysis. SPIS is now used in a much wider scope of applications than initially expected, from mission definition to in-flight scientific measurements analysis. The first version of a new modernized generation SPIS 5.1 aims to address these new issues and challenges in spacecraft surface interaction modeling, including spacecraft charging of commercial missions in Medium Earth Orbit/GEO. For this, SPIS 5.1 offers a simplified user interface and introduces new state-of-the-art software technologies.

Simulation and Analysis of Spacecraft Charging Using SPIS and NASCAP/GEO

Spacecraft charging in GEO particularly concerns dielectric surfaces that may charge to significant voltages relative to spacecraft ground because of the space environment. Testing materials helps to define the level of risk and to maintain confidence in a spacecraft's immunity to damaging effects. Another factor defining the risk involves numerical simulation of spacecraft charging. Several tools aim to calculate surface charging, which is particularly hazardous in harsh environments produced by geomagnetic sub storms, where particles in the energy range of a few to hundreds of kiloelectronvolts are present. The main codes include Nascap-2k, Spacecraft plasma Interaction Software (SPIS), MUSCAT, and Coulomb-2. They use different numerical and sometimes physical models and cross checking their results is a necessary process to achieve better confidence in simulations performed by spacecraft prime manufacturers. The objective of this paper is to simulate different GEO spacecraft configurations with NASA Charging Analyzer Program at geosynchronous orbits (a 1980s to 1990s predecessor to Nascap-2k) and SPIS and to compare the results, both in terms of absolute and differential potentials. The first section concerns the SCATHA spacecraft. The second part of this paper compares efforts to model a modern telecom spacecraft. Finally, we conclude on the reliability of the simulations performed and possible areas for modeling improvement.

New SPIS Capabilities to Simulate Dust Electrostatic Charging, Transport, and Contamination of Lunar Probes

The spacecraft–plasma interaction simulator has been improved to allow for the simulation of lunar and asteroid dust emission, transport, deposition, and interaction with a spacecraft on or close to the lunar surface. The physics of dust charging and of the forces that they are subject to has been carefully implemented in the code. It is both a tool to address the risks faced by lunar probes on the surface and a tool to study the dust transport physics. We hereby present the details of the physics that has been implemented in the code as well as the interface improvements that allow for a user-friendly insertion of the lunar topology and of the lander in the simulation domain. A realistic case is presented that highlights the capabilities of the code as well as some general results about the interaction between a probe and a dusty environment.

An Update of Spacecraft Charging Research in India: Spacecraft Plasma Interaction Experiments—SPIX–II

Arcing on spacecraft solar panels has been a major challenge since it was first observed in late 1970s. Scientists of leading scientific countries like USA, Japan, France, and others are working hard to understand various aspects of this charging-arcing phenomenon. Based upon the experimental simulation results, a few arc mitigation techniques have been developed. In 2003, the Indian Space Research Organization (ISRO) realized the necessity to develop an experimental setup for understanding the charging-arcing phenomenon on satellite's solar panels when parked in lower earth orbit (LEO) in the Space Plasma Interaction Experiments (SPIX) facility. The SPIX facility was jointly developed as a collaborative research project between the ISRO and the Institute for Plasma Research. In 2010, an initiative was taken for the upgradation of the SPIX facility for testing both LEO and geosynchronous earth orbit (GEO) satellite's solar panels by the following standard ISO 11221 experimental procedure and indigenously developed high-speed data acquisition and analysis system. Various types of solar panel coupons were tested in upgraded SPIX-II facility. Sample solar panel coupons used in the experiments can be categorized based on the types of solar cells and the configurations of solar cell that includes different interstring gaps and grouting. For LEO satellite coupons, depending upon the types of solar panel coupons, different bias voltages were applied and threshold voltage was determined. Threshold arc tests were done at different bias voltages and electron gun energies for GEO satellite's coupons. Arc rates and threshold voltages were measured for each type of solar panel coupon, which determined their durability and reliability in harsh environmental conditions. Grouted solar panel coupons had higher threshold voltage in both LEO and GEO worst environmental conditions. Results showed good repeatability and reproducibility for both LEO and GEO experiments. To measure sustained arc conditions, experiments were performed with different string voltages and string currents in both LEO and GEO environmental conditions. It was experimentally validated that the probability of temporary sustained arc or permanently sustained arc increases on increasing either string voltage or string current. In addition, an attempt is also made to understand the role of plasma curing of solar panel coupons and its impact on arc mitigation for GEO environment like space conditions.

SPIS 5: New Modeling Capabilities and Methods for Scientific Missions

Since the last version, the numerical core and the user interface of Spacecraft Plasma Interaction Software (SPIS) have been significantly improved to achieve two objectives: 1) to make SPIS more user friendly and robust for industrial use and 2) to extend the multiscale capabilities and the precision of the solvers in order to model a large range of scientific missions. The new numerical algorithm and modeling capabilities are presented in detail. This new version permits modeling of time variations of the plasma environment, spinning spacecraft, semitransparent grids, secondary emission from 1-D thin elements (e.g., wires or booms), 2-D thin elements (for example, solar arrays), the effect of $v\times B$ electric field, particle detectors, and Langmuir probes onboard spacecraft.

Characterisation of charging kinetics of dielectrics under continuous electron irradiation through real time electron emission collecting method

Dielectric materials used for spacecraft applications are often characterised under electron irradiation in order to study their physical and electrical mechanisms. For surface potential measurement, a small removable flat device based on the secondary electron spectrometer method has been developed and installed in the CEDRE irradiation test facility at ONERA (Toulouse, France). This technique was developed to get rid off specific issues inherent to the Kelvin Probe technique. This experimental device named REPA (Repulsive Electron Potential Analyser) allows in situ and real time assessment of the surface potential built up on dielectric materials under continuous electron irradiation. A calibration has been performed in order to validate this experimental setup. Furthermore, to optimise its efficiency, the physical behaviour of this device has been modelled and numerically simulated using Particle In Cell (PIC) model and a dedicated numerical code called SPIS (Spacecraft Plasma Interactions System). In a final step, electrical characterisations of a charged dielectric have been carried out under continuous electron irradiation with this new method. These results have been compared with measurements performed in same experimental conditions with conventional Kelvin Probe method. The experimental results have been discussed in this paper. To conclude, advantages of this experimental setup in regard of this application will be emphasised.

An Overview of Spacecraft Charging Research in India: Spacecraft Plasma Interaction Experiments—SPIX-II

Arcing on spacecraft solar panels has been a major challenge since it was first observed in late 1970s. Scientists of leading scientific countries like USA, Japan, France, and others are working hard to understand various aspects of this charging–arcing phenomenon. Based upon the experimental simulation results, a few arc mitigation techniques have been developed. In 2003, the Indian Space Research Organization (ISRO) realized the necessity to develop an experimental setup for understanding the charging–arcing phenomenon on satellite’s solar panels when parked in lower earth orbit (LEO) in the Space Plasma Interaction Experiments (SPIX) facility. The SPIX facility was jointly developed as a collaborative research project between the ISRO and the Institute for Plasma Research. In 2010, an initiative was taken for the upgradation of the SPIX facility for testing both LEO and geosynchronous earth orbit (GEO) satellite’s solar panels by the following standard ISO 11221 experimental procedure and indigenously developed high-speed data acquisition and analysis system. Various types of solar panel coupons were tested in upgraded SPIX-II facility. Sample solar panel coupons used in the experiments can be categorized based on the types of solar cells and the configurations of solar cell that includes different interstring gaps and grouting. For LEO satellite coupons, depending upon the types of solar panel coupons, different bias voltages were applied and threshold voltage was determined. Threshold arc tests were done at different bias voltages and electron gun energies for GEO satellite’s coupons. Arc rates and threshold voltages were measured for each type of solar panel coupon, which determined their durability and reliability in harsh environmental conditions. Grouted solar panel coupons had higher threshold voltage in both LEO and GEO worst environmental conditions. Results showed good repeatability and reproducibility for both LEO and GEO experiments. To measure sustained arc conditions, experiments were performed with different string voltages and string currents in both LEO and GEO environmental conditions. It was experimentally validated that the probability of temporary sustained arc or permanently sustained arc increases on increasing either string voltage or string current. In addition, an attempt is also made to understand the role of plasma curing of solar panel coupons and its impact on arc mitigation for GEO environment like space conditions.

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.

LEOrbit: A program to calculate parameters relevant to modeling Low Earth Orbit spacecraft–plasma interaction

Abstract Understanding the physics of interaction between satellites and the space environment is essential in planning and exploiting space missions. Several computer models have been developed over the years to study this interaction. In all cases, simulations are carried out in the reference frame of the spacecraft and effects such as charging, the formation of electrostatic sheaths and wakes are calculated for given conditions of the space environment. In this paper we present a program used to compute magnetic fields and a number of space plasma and space environment parameters relevant to Low Earth Orbits (LEO) spacecraft–plasma interaction modeling. Magnetic fields are obtained from the International Geophysical Reference Field (IGRF) and plasma parameters are obtained from the International Reference Ionosphere (IRI) model. All parameters are computed in the spacecraft frame of reference as a function of its six Keplerian elements. They are presented in a format that can be used directly in most spacecraft–plasma interaction models. Program summary Program title: LEOrbit Catalogue identifier: AENY_v1_0 Program summary URL: http://cpc.cs.qub.ac.uk/summaries/AENY_v1_0.html Program obtainable from: CPC Program Library, Queen’s University, Belfast, N. Ireland Licensing provisions: Standard CPC licence, http://cpc.cs.qub.ac.uk/licence/licence.html No. of lines in distributed program, including test data, etc.: 270308 No. of bytes in distributed program, including test data, etc.: 2323222 Distribution format: tar.gz Programming language: FORTRAN 90. Computer: Non specific. Operating system: Non specific. RAM: 7.1 MB Classification: 19, 4.14. External routines: IRI, IGRF (included in the package). Nature of problem: Compute magnetic field components, direction of the sun, sun visibility factor and approximate plasma parameters in the reference frame of a Low Earth Orbit satellite. Solution method: Orbit integration, calls to IGRF and IRI libraries and transformation of coordinates from geocentric to spacecraft frame reference. Restrictions: Low Earth orbits, altitudes between 150 and 2000 km. Running time: Approximately two seconds to parameterize a full orbit with 1000 points.

Solar wind plasma interaction with solar probe plus spacecraft

Abstract. 3-D PIC (Particle In Cell) simulations of spacecraft-plasma interactions in the solar wind context of the Solar Probe Plus mission are presented. The SPIS software is used to simulate a simplified probe in the near-Sun environment (at a distance of 0.044 AU or 9.5 RS from the Sun surface). We begin this study with a cross comparison of SPIS with another PIC code, aiming at providing the static potential structure surrounding a spacecraft in a high photoelectron environment. This paper presents then a sensitivity study using generic SPIS capabilities, investigating the role of some physical phenomena and numerical models. It confirms that in the near- sun environment, the Solar Probe Plus spacecraft would rather be negatively charged, despite the high yield of photoemission. This negative potential is explained through the dense sheath of photoelectrons and secondary electrons both emitted with low energies (2–3 eV). Due to this low energy of emission, these particles are not ejected at an infinite distance of the spacecraft and would rather surround it. As involved densities of photoelectrons can reach 106 cm−3 (compared to ambient ions and electrons densities of about 7 × 103 cm−3), those populations affect the surrounding plasma potential generating potential barriers for low energy electrons, leading to high recollection. This charging could interfere with the low energy (up to a few tens of eV) plasma sensors and particle detectors, by biasing the particle distribution functions measured by the instruments. Moreover, if the spacecraft charges to large negative potentials, the problem will be more severe as low energy electrons will not be seen at all. The importance of the modelling requirements in terms of precise prediction of spacecraft potential is also discussed.

SPIS and MUSCAT Software Comparison on LEO-Like Environment

Two spacecraft charging software tools, the Spacecraft Plasma Interaction Software and the Multi-Utility Spacecraft Charging Analysis Tool, have been compared in the situation of a wake generated by an object immersed in dense drifting plasma. Each model takes account of the particle dynamics, the space charge effect on the electric field, and the currents collected by the object. The cross-comparison resulted in a good agreement between the two codes and with previous experiments conducted on the same configuration. In particular, a highly negative voltage imposed to the object rear side allows the collection of drifting ions at this location. The nontrivial shape of the current density map was correctly simulated by the two codes. Future works may allow to reach a better quantitative agreement.

SPIS Multitimescale and Multiphysics Capabilities: Development and Application to GEO Charging and Flashover Modeling

While it demonstrated its capability to simulate many common situations, Spacecraft Plasma Interaction Software (SPIS) open source code lacked the possibility to model more challenging situations. The major cases of interest were identified as related to either multitimescale or multiphysical situations. Two major improvements were brought to SPIS code to answer these needs. The first one was an implicit circuit solver with an automatic determination of time steps. The implemented Newton-type algorithm makes use of a predictor for the plasma current variations when surface potentials change. The second major development was a multiphysical model for electrons. The approach consisted in using equilibrium or dynamical electron models in two different zones and connecting these zones at their boundary through a Child-Langmuir (CL)-type condition. Fulfilling this condition dynamically determines the location of the boundary and the electron current to be injected from the dense thermal zone to the space-charge zone. These new features were then used to simulate validation and application cases. The first one consisted in modeling a charging situation in GEO, which had been modeled with NASA charging analyzer (NASCAP) codes and published. Concerning the multiphysical model of electrons, the loop controlling the CL condition was first tested on elementary cases, exhibiting a good qualitative and quantitative behavior. It was then applied to the modeling of a ground experiment performed in JONAS plasma tank at the Office National d'Etudes et Recherches Aerospatiales, i.e., the expansion of the flashover generated by an electrostatic discharge over a precharged solar array coupon, leading to its neutralization.

Simulation of Potential Measurements Around a Photoemitting Spacecraft in a Flowing Plasma

Plasma measurements by electrostatic probes are influenced by the spacecraft-plasma interaction, including the photoelectrons emitted by the spacecraft. Such effects get particularly important in tenuous plasmas with large Debye lengths. We have used the particle-in-cell code package SPIS to study the close environment of the Rosetta spacecraft, and the impact of the spacecraft-plasma interaction on the electrostatic potential at the position of the Langmuir probes onboard. The simulations show that in the solar wind, photoemission has a bigger impact than wake formation. Spacecraft potential estimates based on Langmuir probe data in the solar wind need to be compensated for these effects when the spacecraft attitude varies. The SPIS simulations are validated by comparison to an independent code.

Modeling of Cassini's charging at Saturn orbit insertion flyby

[1] Spacecraft-plasma interactions are studied in the regime relevant for Cassini during the Saturn orbit insertion (SOI) flyby. Three-dimensional particle-in-cell self-consistent code has been employed to calculate the spacecraft potential in a complicated plasma environment of Saturn's magnetosphere for a wide range of distances (for dipole shells between L ∼ 4 and L ∼ 10). Plasma parameters derived from Cassini plasma spectrometer (CAPS) measurements at the SOI flyby have been used as input data. Modeling includes photoemission due to the solar UV radiation, two populations of ions, plasma flows and the flyby geometry. Numerical results highlight significant differences in the plasma conditions along the SOI trajectory. Outside Rhea's shell, at L ∼ 10, the positive spacecraft potential is primary due to the photoelectron production on the Sun-exposed surface of the orbiter. In the inner plasmasphere (L < 8) the spacecraft charges to a negative potential with values up to a few electron temperatures. The dependence of the potential versus distance from Saturn demonstrates behavior qualitatively similar to that deduced from the CAPS data acquired during the SOI period. Simulations display different spatial structures of the plasma particle densities and the self-consistent potential surrounding the SC at different distances from Saturn. Three plasma constituents (electrons, water group ions and protons) produce a new kind of plasma wake with a self-consistent separation between the plasma components in the electric field of the orbiter. Ion focusing occurs in the inner magnetosphere only for protons while the heavy water group ions always form a geometric wake downstream of the spacecraft. The main factor defining the spatial configurations of the potential and of the plasma density perturbations is the energy distribution of the surrounding plasmas. The obtained results can be of importance for understanding the main physical processes occurring in Saturn's magnetosphere, including dust-plasma interactions in the planetary rings, and for reliable interpretations of the electric field and plasma parameter measurements.

Numerical Simulation of SMART-1 Hall-Thruster Plasma Interactions

SMART-1 has been the first European mission using a Hall thruster to reach the moon. An onboard plasma diagnostic package allowed a detailed characterization of the thruster exhaust plasma and its interactions with the spacecraft. Analysis of in-flight data revealed, amongst others, an unpredicted large cyclic variation of the spacecraft floating potential and a mysterious asymmetry in the plasma surrounding the spacecraft. To investigate the details of the anomalies, we developed the numerical software tool SmartPIC to characterize and predict spacecraft-plasma interactions. Technical details, such as solar arrays and onboard diagnostic devices, have been modeled with high accuracy. All basic plasma parameters, the spacecraft floating potential, backflow distributions, and ion impact energies are calculated by the code and are available in high spatial resolution throughout the computational domain containing the entire satellite. It was possible to clearly identify the rotating solar cells arrays as the source of the cyclic variation of the spacecraft floating potential. Furthermore, the asymmetry of the plasma formation around the spacecraft is linked to the location of the neutralizer causing a region of increased charge-exchange collisions particle generation. Both of these results have significant impact on the implementation of electric propulsion systems on satellites.

New electromagnetic particle simulation code for the analysis of spacecraft-plasma interactions

A novel particle simulation code, the electromagnetic spacecraft environment simulator (EMSES), has been developed for the self-consistent analysis of spacecraft-plasma interactions on the full electromagnetic (EM) basis. EMSES includes several boundary treatments carefully coded for both longitudinal and transverse electric fields to satisfy perfect conductive surface conditions. For the longitudinal component, the following are considered: (1) the surface charge accumulation caused by impinging or emitted particles and (2) the surface charge redistribution, such that the surface becomes an equipotential. For item (1), a special treatment has been adopted for the current density calculated around the spacecraft surface, so that the charge accumulation occurs exactly on the surface. As a result, (1) is realized automatically in the updates of the charge density and the electric field through the current density. Item (2) is achieved by applying the capacity matrix method. Meanwhile, the transverse electric field is simply set to zero for components defined inside and tangential to the spacecraft surfaces. This paper also presents the validation of EMSES by performing test simulations for spacecraft charging and peculiar EM wave modes in a plasma sheath.

Evaluation of Electric Field Probe Onboard Spacecraft Using a 3D Full PIC Simulation

A three-dimensional electrostatic full Particle-In-Cell code has been developed to analyze spacecraft-plasma interactions quantitatively. The code is expected to achieve highly accurate computation because it has no robust algorithm. We adopted the code to evaluate the electric field measurement onboard spacecraft, especially under low-density ambient plasma environment with photoelectron emission. In this paper, first, fundamental functions for computation of the electric field were validated in Low Earth Orbit and Geosynchronous Orbit environments by comparing with the thin-sheath limit theory and the thick-sheath limit theory, respectively. Second, the floating potential of a spacecraft model with photoelectron emission in Geosynchronous Orbit was computed to examine the effects of photoelectron emission to the electric potential around the spacecraft. The dependence of the floating potential on the photoelectron temperature was shown in the simulation.

Estimation of Auroral Environment by Electrostatic Full-particle Simulations Modeling of REIMEI Satellite Observations

A full particle-in-cell (PIC) plasma simulator with higher accuracy as well as reasonable performance for studying spacecraft-plasma interactions has been developed. It is applied to estimate characteristics of the onboard current monitors (CRM) for a small scientific satellite REIMEI in the polar orbit, because calibration is required for the observation data. Basic characteristics in typical plasma environments including effects of the satellite geometry are demonstrated.

Accomplishment of multi-utility spacecraft charging analysis tool (MUSCAT) and its future evolution

Abstract (MUSCAT) is a high value computation tool for analyzing spacecraft–plasma interaction, whose typical example is charging–arcing issue, corresponding to spacecrafts in LEO, GEO and PEO. JAXA and Kyushu Institute of Technology (KIT) started the development as a joint project in November 2004 and the final version of MUSCAT was released in March 2007. The final version includes many important features to simulate spacecraft–plasma interaction and the features can be separated into four parts. The first part is its GUI named “Vineyard”. By using Vineyard, MUSCAT users can build a satellite model including not only its geometry but also material properties of the surface. As for the second part, MUSCAT includes many kinds of effects derived from space plasma environment as well as electrical functions of spacecraft. For the third part, MUSCAT can work on parallel workstation with multi-CPU. The last feature is that the computation result by MUSCAT was thoroughly validated by experiments in plasma chamber. The numerical result shows very good agreement with the code validation experiment. We also conducted trial computation of charging analysis on Greenhouse gases Observing Satellite (GOSAT) with MUSCAT. One purpose of the computation was prediction of charging status of GOSAT for the real satellite design in combination with the ground test. The other is performance assessment of MUSCAT. After the joint project, expansion and maintenance of MUSCAT will be carried out by “MUSCAT Space Engineering Ltd” which is a new enterprise made of the development team. In future we will try to conduct MUSCAT computation for various spacecrafts and also try to add useful function such as 3D CAD compatibility.