16th Spacecraft Charging Technology Conference

Physical Sciences Inc. (PSI) has developed a paintable coating that can be applied directly to spacecraft and spacecraft components or pre-applied to laminar materials for installation on spacecraft. The paintable coating will provide spacecraft operating at altitudes between low Earth orbit to geosynchronous orbit, an added layer of protection against catastrophic electrostatic discharge events. Those events can result in a potential total loss scenario, that with this technology can easily and affordably be guarded against. The coating is passive - no electrical wiring connection to or power from the spacecraft is required. The coating developed under this program has negligible impact on the size, weight, power and cost (SWaP-C) of the spacecraft, decreasing risk to assets at a low premium investment. The benefits of the coating to commercial satellites are the same as for the military and scientific missions’ applications – protection of expensive, mission critical, difficult to replace space assets with minimal cost, effort and impact on the mission. Additional benefits of this research (and the solid-state cathode devices generated in addition to the coatings) include potential applications to electrodynamic tether propulsion and terrestrial spinoff applications in the realm of powered cathode devices.

Jonathan Rameau

Physical Sciences Inc.

As Langmuir probes (LP) transition to small spacecraft (S/C), such as the increasingly popular CubeSats, their implementation encounters new challenges that aren’t present on larger spacecraft. LP operation on CubeSats impact S/C charging levels due to the limited ion-current collection to the spacecraft body relative to the electron current collection by the Langmuir probe. This variable, net-negative S/C potential will introduce errors in temperature and density measurements limiting the effectiveness of LP measurements [1]. One method of counteracting the effects of a variable spacecraft potential is to track the changes in spacecraft charging using a high impedance probe during LP operations, the twin-probe method. Experiments performed at NASA Marshall Space Flight Center demonstrated the feasibility of the twin-probe method as the SC to LP surface area ratio decreases [2].

These experiments were a beneficial first step to validating the twin-probe method, however there remained unanswered questions due to chamber effects, such as charge exchange ions and high background neutral pressures. Understandably, performing in-chamber and in-orbit experiments for all possible spacecraft and instrument configurations is infeasible and prohibitively expensive. To predict the spacecraft’s charging behavior, MATLAB software, called the Plasma Spacecraft Interaction Codes (PSIC), was developed. The PSIC perform ”back of the envelope” calculations to estimate spacecraft charging behavior and understand how Langmuir probe measurements are impacted by a variable spacecraft potential. Here we present the general results of the code and comparisons to chamber measurements that highlight good agreement between models and chamber measurements. Then we will discuss the effectiveness of the twin-probe method over a range of parameters such as area ratio and the input impedance of the high impedance probe. Finally, we will present additional applications that are feasible with a twin-probe system on a CubeSat sized platform, such as maintain a negative CubeSat potential to enhance ion energy analyzer performance.

[1] L. H. Brace, ""Langmuir probe measurements in the ionosphere,"" Measurement Techniques in Space Plasmas: Particles, pp. 23-25, 1998. [2] O. Leon, J. McTernan, J. A. Vaughn, T. A. Schneider, G. Miars, W. Hoegy, B. Gilchrist, ”The Twin-Probe Method: Improving Langmuir probe measurements on small spacecraft,” IEEE Transactions on Plasma Science. January, 2022. doi: 10.1109/TPS.2021.3137765.

Omar Leon

University of Michigan

Space is a particularly brutal environment and especially so for polymers which can be exposed to ionizing radiations (electrons, protons, photons), ultra-violet rays, atomic oxygen, extreme temperature variations and electrostatic discharge (ESD) issues.

This paper presents some of the latest results concerning the development of new range of anti-static wires and cables solutions with controlled volume resistivity in order to mitigate surface and internal electrostatic discharges (ESD). This activity was initiated and funded by ESA in the scope of ESA ITT AO/1-7877/14/NL/RA.

■ NEW ESD-RESISTANT ANTISTATIC WIRES INSULATIONS

□ Electrostatic Discharge issue in Space

Wires and cables used in space can be susceptible to initiating electrostatic discharges (ESD) under certain conditions. The origin of ESD events is mainly due to the presence of excellent electrical insulation layers and their ability to store and build-up electrostatic charges due to their intrinsic high resistivity.The incoming charged particles can build-up inside dielectrics to “break down” threshold levels leading to electrical arcs discharged into nearby sensitive circuits.

□ New ESD-resistant antistatic wires concept

A standard way to protect wires and cables and others ESD sensitive components is to cover them with conductive films, braids or coatings in order to absorb and evacuate most of incoming charged particles toward the mass of the satellite. The drawback of this solution is the increase in mass, cost and volume of the finished product, and reduced flexibility. The thickness of any additional conductive material also needs to be optimized in order to reduce the residual radiation flux enough to fully prevent any internal ESD (iESD) risks. The added conductive layer also has to be grounded which can further increase the mass, cost and complexity of the systems.

An innovative approach explored by Axon’ to reduce the ESD risk for wires and cables without these drawbacks is to fully replace all the highly insulating dielectrics with “leaky” materials in order to permit a fast enough evacuation of incoming charged particles to the nearest conductive layer. The main objective of this development, supported by ESA activity ITT AO/1-7877/14/NL/RA, was to explore this concept by formulating, processing and testing new wire and cable insulation materials able to dissipate through to the central conductor the electrostatic charges that build up in a space environment.

The main challenge was to formulate and process new materials with volume resistivities low enough to permit sufficient charge decay speed (for Space-ESD resistance), but at the same time maintaining sufficiently high values to guarantee the electrical insulation performance of the finished wires (acceptable insulation resistance and dielectric strength).

□ Development Results

New materials and AWG 24 coated antistatic wires prototypes with volume resistivities in the range of ≈1E+7 to 1E+15 ohm.cm have been successfully manufactured and characterized.

Electrical, mechanical and thermal tests have been performed based on ESCC3901 specifications, like spark test, voltage test, abrasion, cut-through, blocking, shrinkage, cold bend and thermal ageing.

Complementary tests have also been performed such as tensile properties, breakdown voltage, dielectric strength, and volume resistivity measurements as function of temperature and applied voltage.

Pierre-Yves MIKUS

Axon Cable

Spacecraft charging, defined as the buildup of charge in and on spacecraft materials, is a significant phenomenon for spacecraft in certain Earth and other planetary environments. Space charging effects are caused by interactions of the in-flight plasma environment (mostly electron) with spacecraft materials and electronic subsystems. In fact, surface charging/discharging and Internal Electrostatic Discharging (IESD) have been identified as the primary causes of spacecraft anomalies and failures [1]. Possible detrimental effects of spacecraft charging include disruption of or damage to subsystems (power, navigation, communications, instrumentation, etc.) because of charge buildup and electrostatic discharge. Charges can also attract contaminants, affecting thermal properties, optical instruments, and solar arrays, and can change particle trajectories, thus affecting plasma-measuring instruments.

Design for control and mitigation of surface charging, the buildup of charge on the exterior surfaces of a spacecraft due to space plasmas, is treated in detail in NASA TP2361, Design Guidelines for Assessing and Controlling Spacecraft Charging Effects (September 1984). To address the growing concerns at the time associated with the in-flight buildup of charge on internal spacecraft components due to space plasmas with high energy electrons, NASA-HDBK-4002, Avoiding Problems Caused by Spacecraft On-Orbit Internal Charging Effects, was written by Henry Garrett and Albert Whittlesey in 1999 as a companion document to NASA TP2361. Although many of the ideas presented in NASA-HDBK-4002 had a long heritage, NASA-HDBK-4002 collected them in one convenient place and quantified and illustrated the design guidelines necessary to reduce the internal charging effects for the first time.

Since the original writing of the two documents, there had been developments in the understanding of spacecraft charging issues and mitigation solutions, as well as advanced technologies needing new mitigation solutions. That, and the desire to merge the two documents, was the motivation for the revision, NASA-HDBK-4002A, Mitigating In-space Charging Effects—a Guideline, written by Henry Garrett and Albert Whittlesey in 2011. This revision, the current NASA charging handbook (NASA-HDBK-4002A) contains details of the spacecraft design procedures for minimizing detrimental effects of spacecraft charging and for limiting the effects of the resulting electrostatic discharge. As such, it has served as the primary document for evaluating, testing, and mitigating surface and internal charging effects.

Since the last revision in 2011, however, there have been 4 major spacecraft charging conferences. The last 4 spacecraft charging technology conferences have addressed new levels of electron environments, new modeling methods, and many laboratory tests that are now widely accepted, which need to be incorporated in the charging handbook. The revised NASA charging handbook, NASA-HDBK-4002B, which is being published, will be presented. One thing to be noted is that 4002 and 4002A were expressly intended to be written to be guidelines rather than requirements. Unfortunately, many users merely copy various equations in the Handbook, which can lead to unfortunate results, because the mission and experiment needs can be vastly different for different projects. We clarify this situation for 4002B.

Koons, H. C., J. E. Majur, R. S. Selesnick, J. B. Blake, J. F. Fennell, J. L. Roeder, and P. C. Anderson, “The Impact of the Space Environment on Space Systems,” Proceedings of the 6th Spacecraft Charging Technology Conference, September, 2000.

Wousik Kim

Jet Propulsion Laboratory, California Institute of Technology

The electron-heavy Jovian environment is challenging for the science missions that explore Jupiter and its moons. Missions such as Juno, JUICE, and NASA’s Europa Clipper mission contend with very low temperatures and a severe radiation environment that must be evaluated for effects including dose, dose rate, and Internal Electrostatic Discharge (IESD) effects.

The use of carbon fiber composites is ubiquitous within the modern spacecraft industry owing to its desirable mechanical, thermal, and RF properties. It is known, however, that these materials are susceptible to spacecraft charging and electrostatic discharges (ESD) as a result of non-conductive surface resins present in the carbon fiber composite layup. The Europa Clipper High Gain Antenna (HGA) represents a large (~3 m diameter) carbon fiber composite assembly present on the Clipper spacecraft. Multiple combinations of fiber and resin are present throughout the reflector, struts, and support systems. Owing to its large size and prominence within the Clipper Flight System (FS), discharges on / in the HGA could impact both FS RF communications and science payload operations – specifically the REASON (Radar for Europa Assessment and Sounding: Ocean to Near-surface) instrument. Severity of such impacts may range from “catastrophic” (e.g. overstress and permanent destruction of the victim electronics) to “nuisance” (e.g. noisy communications or degraded science observations). Regardless of severity ESD poses a threat not only to spacecraft systems but also meeting science / mission requirements.

In vacuo flood beam or similar test conditions are necessary to properly characterize the charging and IESD hazards as familiar, contacting, electrical measurements are overly influenced by the conductive carbon fibers. The effects of Radiation Induced Conductivity (RIC) on the charging and ESD hazard – specifically on discharge magnitude and rate – may be significant. Simultaneous inclusion of a second electron beam enables better matching of the predicted in-flight charge deposition rate, energy deposition rate, and ultimately less conservative and better representative results. In [1] we discussed the motivation for, and process of, multiple electron beam ground testing. Here we share complete results for all materials tested along with a detailed summary of the recently complete qualification effort of the BR127NC-ESD static dissipative primer used to mitigate charging / discharging on the reflector surface.

A series of tests and analyses were performed between 2020 and 2022 to assess the ESD threat resulting from Clipper spacecraft passage though the Jovian energetic particle regions. The results summarized herein focuses, primarily, on the ESD threat associated with electron charging at End Of Life (inclusion of both cryogenic temperatures and mission Total Ionizing Dose (TID)). This manuscript summarizes test details, methods, and results for a series (>10) of experiments performed by the Johns Hopkins University Applied Physics Laboratory (JHU APL), the NASA Jet Propulsion Laboratory (JPL), and Lockheed Martin Advanced Technology Center (ATC). We also summarize limitations associated with ground testing and extension of results to developing a fully qualified system for the Clipper mission.

References: Likar, J., et al. “Spacecraft Charging and IESD Characterization of Carbon Composite Materials with Multiple Electron Beams.” Presented at Applied Space Environments Conference (ASEC), 1-5 November 2021.

Justin Likar <justin.likar@jhuapl.edu>

The Johns Hopkins University Applied Physics Laboratory, NASA Jet Propulsion Laboratory, Lockheed Martin Advanced Technology Center, Applied Aerospace Structures Corporation, and The Pennsylvania State University

A novel system was developed to simulate the combined effects of reduced gravity and ionizing radiation present during spaceflight on biological and particulate samples. The miniature rotary cell culture system (mRCCS) was designed to synchronously rotate up to five independent vessels containing particulate samples suspended in fluid media, constructed using radiation tolerant, biocompatible, and vacuum compatible materials. Reduced gravity conditions were achieved when suspended particles (e.g., 200 μm polystyrene microcarrier beads with or without adhered cell clusters) were suspended inside the vessels moving near terminal velocity in viscous neutral-buoyant fluid media with densities matched to the suspended particles to achieve neutral buoyancy. Variations in centripetal acceleration from slow rotation of the vessels limited reduced gravity environments from ~1·10^-5 to ~2·10^^-2 g, comparable to similar commercially available systems. The effective gravitational acceleration applied to particles was calibrated through particle tracking of suspended particles within the mRCCS systems vessels. The entire mRCCS apparatus can be used in a standalone configuration for independent reduced gravity simulations or can be introduced into the Utah State University’s Space Survivability Test (SST) chamber for radiation exposure or simultaneous radiation exposure under reduced gravity. The SST chamber has a ~90 mCi 90Sr source that emits 0.2 to 2.5 MeV β radiation. The combined mRCCS and SST chamber system can provide average effective dose rates for the suspended particles, controlled over a broad range (>900X) from ~3.7 mGy/day to 3.4 Gy/day by varying the source-to-sample distance and using varying slit width graphite shields. This system can provide stable, simultaneous space-like radiation and reduced gravity environments for experiments conducted on timescales of minutes to months. Initial experiments have focused on understanding cellular damage due to the effects of radiation and reduced gravity on cardio and neurological cell clusters, with a long-term goal of studying damage mitigation of biological reagents. However, an environment where long-term gravitational forces can be minimized allows the study of electrostatic interactions between particles with very small internal charges (induced perhaps from electron or photo yields), in experiments similar to the seminal Millikan oil drop experiment.

Achal Duhoon

Utah State University

Spacecraft surface charging is induced by the interaction of ambient plasma with the spacecraft surface and sometimes causes spacecraft anomalies due to electrostatic discharging (ESD) in the medium earth orbit (MEO). We analyze the surface charging environment in the MEO using the Helium Oxygen Proton Electron (HOPE) data of the Van Allen Probes from 2012 to 2019. The severe charging (< -1 kV) events are observed only in the eclipse region due to the lack of the photoelectron emission mitigating negative charging and are induced by enhancement of the hot electron flux above 4 Earth radii. These events show good correlations with geomagnetic activity indicated by the Kp and AE indices. The Kp is the index of the average geomagnetic disturbance in the entire magnetosphere. The AE is the index of the auroral electrojet currents and shows the substorm activity such as the hot plasma injection into the inner magnetosphere from the geomagnetotail. The results confirm that the severe charging is induced by the hot plasma injection into the MEO at geomagnetic substorm. The increase of the high energy (10-50 keV) integral electron flux induces the severe surface charging events. In the two-thirds of the severe charging paths, the low energy (0.1~2 keV) and middle energy (2~10 keV) integral electron flux has increased lager than their average levels before the surface charging occurs. That may suggest that the active geomagnetic features from the growth phase to onset of magnetospheric substorm in the MEO. However, in the one-fifth of the severe charging paths, the middle and the high energy electron flux has suddenly enhanced at the same time without the pre-increase of the low and middle energy electron flux using the single spacecraft observation. During the severe surface charging, the ambient low energy electrons are repelled by the negative spacecraft potential and the observed flux of the low energy electrons decreases.

Masao Nakamura

Osaka Metropolitan University

The space territory has complex and dynamic arrangement that can lead to damage the interior and surface of the spacecraft especially at geosynchronous altitude. The Spacecraft (SC) charging threatens various components of spacecraft. Electrostatic charging and mainly discharge can affect the spacecraft surface adversely; the level of spacecraft charging is dependent on particle energy distribution. The present work demonstrates the threshold condition and critical temperature for the onset of significant spacecraft charging. The charging currents due to secondary and back scattered electrons and hence the current balance equation has been developed theoretically by using non-Maxwellian distribution function to refine the correlation between spacecraft and its interactivity with the harsh environment for various space-grade materials. The threshold condition specifies the temperatures and density ratio of both plasma components modifies the current balance equation significantly.

Nazish Rubab <nazish.rubab@ucp.edu.pk>

University of Central Punjab, Pakistan

Space Electrostatic Discharge Effect (SESD) and Single Event Effect (SEE) are two major space environment factors that cause spacecraft failure. Previous studies have established that both of them can cause memory cells to flip. The test in a low-power asynchronous timing monolithic SRAM for satellites was carried out using an ESD generator and a pulsed laser experimental facility. Soft error number characteristics, current value characteristics, and single/multi-bit flip characteristics were compared for similarities and differences. The test revealed that when soft errors occur in the circuits, the current of the power supply drops, allowing the device to be judged by monitoring the current value. It was also discovered that soft errors induced by SEE were mostly single-bit flips, whereas those induced by SESD were mostly multi-bit flips. This study, to some certain extent, provides experimental support for distinguishing device errors caused by these two effects, as well as references for further accurate identification of in-orbit faults and corresponding protection design.

Damon Wang <wangxuan@nssc.ac.cn>

National Space Science Center / University of Chinese Academy of Sciences

Many spacecraft materials are combinations of multiple constituents, sometimes with very different conductivities. Such composite materials include circuit board materials, carbon composites, materials loaded or coated with nanoparticles, etc. Some composites are designed to be leaky dielectrics to avoid the buildup of charge that leads to destructive electrostatic discharge (ESD). Testing in a setup representative of flight conditions is critical to ensure that the desired conduction properties will be present in flight. Composite materials can be sensitive to contact methods, particle concentration, inhomogeneous particle distributions, anisotropy, and significant composition variation among samples of materials with the same trade name. Volume resistivity measurements of dielectric materials loaded with conductive particles are a good example of the difficulties involved since they can be sensitive to the test method used. In a constant voltage parallel plate test configuration, the smallest resistance path—for example, percolation paths through conductive particles—will dominate the measurement and may result in an excessively optimistic high conductivity. The non-contact charge storage method allows charge accumulation in less conductive regions of a material and can potentially result in much higher values of volume resistivity compared to the parallel-plate method. Electron beam-induced ESD tests of carbon composite materials have previously shown discharges. We have tested a version of carbon composite that does not discharge in similar charging conditions. SEM imaging shows significant differences in surface resin distributions compared to carbon composites known to discharge. The variability in charging outcomes from nominally the same materials requires detailed information regarding the origin and composition of the material together with test methods and results. In applications sensitive to spacecraft charging, additional specification beyond material name is needed to ensure that the right version is consistently used. For example, circuit board specifications such as Arlon 85-N or FR-4 offer manufacturers certain flexibility in how boards can be made, and even what materials may be used. We have measured significant differences in the conductivity of circuit board core materials that meet the same board specification under the same name.

Allen Andersen

Jet Propulsion Laboratory, California Institute of Technology

Spacecraft are operated in the radio-active rays environment, such as high energy electrons and protons. The surface of the spacecraft is covered with multi layer insulator (MLI) and optical solar reflector (OSR), which are made of insulating materials. However, these materials are charged up and discharged by high energy electrons and protons. The discharge phenomena are the possible origin of satellite fatal operational anomalies. In general, the surface potential measurement was performed to investigate the charge and discharge characteristics of the charged particles. However, it is difficult to clarify the mechanism of discharge generation because high energy charged particles penetrate in the bulk of material and accumulate. Such accumulated charges form the electric potential at the surface. Therefore, it is important to measure the internal charge in the insulating materials. In this study, we have developed an apparatus that can measure the internal and surface charging of insulating materials under charged particle irradiation using the pulsed electroacoustic (PEA) method (refer to Figure 1). Furthermore, we investigated the reliability of the surface potential and measured the internal and surface charges of fluorinated ethylene propylene (FEP) under electron and proton irradiation (refer to Figure 2 and Figure 3). From the results, the surface potential could be measured with an error of within 6%. When the electron was irradiated, the negative charges were observed in the sample inside and surface, and the amount of their increased with time. Furthermore, the surface potential was -7.5 ~ 13 kV by the accumulated negative charges, then the negative charges were decreased precipitously. The positive charges drifted for the sample irradiation surface and the surface potential was 1.5 kV by the accumulated positive charges. As mentioned earlier, the development of a new apparatus using the PEA method enabled us to measure the internal and surface charging of the insulation material under charged particle irradiation and to calculate the surface potential.

Kaisei Enoki

Tokyo City University

In this paper, we show the characteristics of secondary electron emission on the polyimide degraded by electron.

Spacecraft are operated in an environment with drastic temperature changes. Therefore, it is necessary to maintain a constant temperature inside the spacecraft in order to operate the onboard equipment normally in such an environment, so the surface of the spacecraft is covered by dielectric materials called multi layer insulator (MLI). However, these materials are charged and degraded by high-energy charged particles, such as electron and proton ,degraded during operation. Hence, it is important to consider in achieve a long term operation that charging and discharging phenomena after surface material degraded . Generally, it is assumed that the molecular structure of polymer materials changes when they are irradiated with high-energy charged particles, and actually, it has been reported that molecular chains are scissored by electron beams. Therefore, in this study, we investigated the electron beam-induced secondary electron emission yield (SEEY) of polyimide, which was deteriorated by electron beam irradiation, to examine the changes in the physical properties of the material. As a result, there was a tendency for the SEEY to increase when the dose of degrading electron beam was changed. This research is carried out using the measurement system shown in Fig.1.

Kosuke Amamizu

Tokyo City University

In this paper, we propose to present the charging facilities at ONERA-CNES through the description of different test campaigns and in particular those performed for JUICE mission preparation. A complete charging analysis for a mission such as JUICE requires environment specifications and complete characterization of dielectric materials used for the mission. Surface and bulk conductivity of dielectrics such as coverglasses, polymers or paintings were measured inside SIRENE facility at different temperature levels. SIRENE is a chamber where charging electron environment is reproduced from a few tens of keV to 400 keV, the beam being distributed with a diffusion foils system to simulate GEO orbit or other charging environments and in particular to reproduce both surface charging and radiation induced conductivity processes. Secondary electron emission was also measured on dielectric materials as a function of energy in DEESSE and ALCHIMIE ultrahigh vacuum chambers. In these chambers, secondary electron emission can be characterized in the range of a few eV to 30 keV. The facility has also additional capabilities such as secondary electrons energy analysis and chemical analysis of material surface (XPS, AES), and mass spectrometry. All these tests allowed us to extract the main parameters for a charging analysis by simulation. SPIS simulations could then be performed using these sets of parameters. However, voltage threshold for ESD triggering stays a parameter which is difficult to predict through simulations and the consequences of ESDs on solar cells behavior is obviously not predictable. ESD tests are then necessary to predict ESD voltage threshold, ESD effect on solar cells performance. These tests have been performed at ONERA inside JONAS chamber. This plasma chamber allows us to provide either electrons or plasma for ESD and arcing tests. After design adjustments and modifications, an as-flight coupon for JUICE mission was then tested for ESD and arcing simulation. In addition, solar cells radiation test performed at ONERA on JUICE solar cells will also be evoked with the presentation of radiation test capabilities (AXEL platform with 2 MeV electrons and protons).

Thierry Paulmier

ONERA

Electrostatic discharges (from) electrically isolated conductors (commonly called “floating metal”) were discovered to be sources of anomalies on early spacecraft. Charging and discharging of the ungrounded metallization of multilayer thermal control blankets on DSCS II and Meteosat are examples. Ground tests demonstrated that discharges from floating metal were more severe and hazardous than discharges from equal areas of dielectric film, and led to requirements to ground all thermal blanket layers. Charging and discharging of wires and coaxes were later discovered to be more severe if the conductors were ungrounded, which led to requirements to ground all unused and unterminated wires. In orbit anomalies followed by ground tests also showed that discharges from isolated conductors on printed circuit boards (PCBs) were also a potential risk, despite the much low flux levels inside these units. Ground tests showed that the peak voltage and energy of ESD transients induced in nearby PCB circuits scaled with the area of the area of the floating conductors [Leung, 4th SCTC]. NASA HDBK 4002 (and updated 4002A) recommends limiting the area to 3cm2 (and to 0.3cm2 if floating metal is on a PCB). Recent tests show that the peak voltage and delivered energy depends upon the mode of discharge of the floating metal. Blow-off discharge amplitudes were similar to those observed at JPL for similar areas. But when the discharge punched through the insulator between the floating metal and underlying ground trace, peak voltages and delivered energies were much higher. Given modern designs employ multilayer circuit boards, the opportunity for punch through discharges is higher than existed on the 80mil, single-layer FR4 boards the HDBK 4002 guideline is based upon.

J. Michael Bodeau

Consultant

An analysis of how thick, flat Teflon tape might charge in the GEO and polar LEO space environments was performed with the goal of determining when and if pulsing and/or arcing might occur due to deep electron charging. Multiple simulations were performed using AF-NUMIT3, a dielectric charging simulation model. The polar LEO electron flux environment and the GEO electron flux environment data used in the simulations were based on IRENE for energetic electrons and observational data for auroral electrons. A variety of scenarios were simulated, using environments that were deemed likely to be worst case and average case for charging. For the polar LEO environment, these included both artificially continuous flux and an intermittent set of fluxes designed to approximate the orbit. Simulations were done for a conductor on both sides of the tape and for a conductor only on the backside of the tape. The AF-NUMIT3 simulation generates plots of the electric field developing within the dielectric as a function of depth and time. The simulation showed that if the threshold for danger of pulsing or arcing is accepted to be as low as 10 keV/cm, then in all cases considered for a polar LEO orbit or a GEO orbit there will be significant danger. However, even if the dielectric strength is accepted to be as high as 700 kV/cm, dangerous charging and undoubtedly arcing will certainly occur if there is no conductor on the front surface of the Teflon.

Brian Beecken <beebri@bethel.edu>

Department of Physics and Engineering, Bethel University, St. Paul, MN; Air Force Research Laboratory, Space Vehicles Directorate, Kirtland AFB, NM

Energetic particles in geospace are the most concerned causes of various anomalies from minor errors to critical malfunctions to spacecrafts. One of the energetic particles is radiation belt electrons quasi-permanently trapped in the Earth’s magnetosphere. The energetic electrons cause the internal charging which builds up the electric field inside the dielectric materials. In the case that the electric field exceeds the threshold, the electrostatic discharge (ESD) occurs. The ESD causes material degradation and trouble of electronic equipment, thus the ESD is a concerned threat to the spacecraft.

In order to understand the internal charging and its threats, we developed a numerical model for the internal charging. The model is based on an equivalent circuit of a Resistor and a Capacitor (RC model) to evaluate the temporal variation of the charge accumulation in spacecraft dielectric materials. Electron flux with energies higher than 2 MeV obtained from GOES series in solar cycles 23 and 24 is applied to the RC model to evaluate the internal charging of the material deep inside the spacecraft. Based on the model calculation driven by the GOES data, we found that the charge accumulation tends to be faster during the decline phase of solar cycles, which is consistent with the report by Bodeau et al. (2010). Considering the threshold level of the electric field for the ESD, we incorporated the rapid charge-release effects due to the discharge in the RC model (RCD model). By using the RCD model, we evaluated the temporal variation of the discharge threat for 24 years, and found that the threat of the internal charging in the decline phase of solar cycle 24 was higher than that of solar cycle 23. In this presentation, we report the assessment of the ESD threat of the internal charging of material deep inside spacecraft by the RCD model, using long-term, in-situ data in geostationary orbit.

Shinji Saito

National institute of Information and Communications Technology

The PICASSO (Pico-Satellite for Atmospheric and Space Science Observations) spacecraft, is a 3U BIRA (Belgian Institute for Space Aeronomy)/ESA nanosatellite demonstration mission in Sun Synchronous Low Earth Orbit, hosting in particular the Sweeping Langmuir Probe experiment. It was successfully launched as a passenger payload on a Vega rocket from French Guyana on September 3rd, 2020. The sweeping Langmuir probe instrument (SLP) on PICASSO consists of four thin (1mm radius) cylindrical probes providing measurements of plasma electrons density and temperature, the spacecraft’s floating potential. Simulations were carried out to analyze the behavior of the extracted current-voltage characteristic of the probes as influenced by different ambient plasma parameters but also for different orientations of the spacecraft and solar arrays wrt to the s/c velocity vector, exposing the different surfaces to solar illumination or the wake of the spacecraft [1].

In this study we fully revisit the numerical model of the PICASSO spacecraft, assessing the effects of a refined geometric model including realistic probes sheath representation on the simulated instrument response and benefiting from the latest improvements of the SPIS solvers (SPIS6.1.0), to simulate Langmuir probe sweeps for October 26th, where a number of I-V curves observations are available.

Preliminary comparisons between the observed and simulated I-V curves, as well as interpretation of the observed signals will be provided.

[1] A. Waets, F. Cipriani and S. Ranvier, "LEO Charging of the PICASSO Cubesat and Simulation of the Langmuir Probes Operation," in IEEE Transactions on Plasma Science, vol. 47, no. 8, pp. 3689-3698, Aug. 2019, doi: 10.1109/TPS.2019.2920136.

Jean Porto

ESA

Electrostatic discharge (ESD) is one of the major causes of satellite anomalies. There are two types of charging mechanism that will cause ESD, namely, surface charging caused by keV plasmas and internal charging caused by high-energy electrons. Since space plasmas change with time and space, the plasma environment is different depending on the position of the satellite. In addition, how deeply a satellite is charged by a certain plasma environment depends on the design and material of the satellite. Therefore, the ESD risk is different for individual satellites even if the same plasma environment is given. In order to contribute to safety and stability of satellite operation, we have been developing a charging risk estimation scheme, called SECURES (Space Environment Customized Risk Estimation for Satellites).

As the surface charging part of SECURES, we combine the MHD model and charging models, targeting the GEO region as the first step. Here we had two challenges. (1) The charging analysis tools require the densities and temperatures of ions and electrons (Ni, Ti, Ne, and Te) as the input. Because MHD models does not provide separately the electron and ion parameters, we developed an empirical method to estimate those from MHD data for GEO region. (2) The other problem is that the charging calculation takes a long time, therefore, even if the plasma data is provided in real-time, it is impossible to obtain satellite potentials in real-time. Thus, we have developed a quick estimation method.

Based on the above, we have developed a geosynchronous surface charging assessment system by using the realtime magnetosphere simulation data, and we have upgraded the system. We report the above mentioned fundamental researches and the upgraded system.

Aoi Nakamizo

National institute of Information and Communications Technology

The use of Langmuir probes on small satellites in the ionosphere is problematic due to the induced charging of the spacecraft by the probe. This well-known problem results from a lack of sufficient area on the spacecraft to supply a return current when the probe is collecting electrons. This is further complicated by difficulties in modeling collections of conducting surfaces at various exposed potentials on the spacecraft surface for the meso-thermal plasma conditions on orbit. Utah State University has used a circuit model approach to describe the interaction between various collectors on a small satellite [3]. The I-V relationship for each collector is modeled using custom circuit elements based on orbit motion limited theory for stationary and drifting plasmas [1,2]. The circuit network is then solved using SPICE [4] to understand how spacecraft charging and extraneous exposed potentials effects Langmuir Probe measurements. Other modeling software exists for understanding spacecraft charging, such as NASCAP, which makes use of a 3D model of the spacecraft and performs a particle in cell type simulation. These models have been shown to be very accurate at matching observations and are useful for predicting the physical interactions of a spacecraft body with the space environment; however, they are limited in their capacity to model the complex electrical interactions of the spacecraft analog circuits and instrumentation with the space plasma environment. In this paper we compare the SPICE analytical vs NASCAP for modeling of the SPORT spacecraft, a 6U mission with a Langmuir probe, two floating potential probes, and exposed potentials from a retarding potential analyzer.

REFERENCES Mott-Smith, H. M. and I. Langmuir, “The Theory of Collectors in Gaseous Discharges”, (1926), Phys. Rev., 28, 4. https://link.aps.org/doi/10.1103/PhysRev.28.727

Hoegy, W.R. and L.E. Wharton, “Current to moving spherical and cylindrical electrostatic probes.” 1971. Barjatya, A., ""Langmuir Probe Measurements In The Ionosphere"" (2007). All Graduate Theses and Dissertations. 274. https://digitalcommons.usu.edu/etd/274 Quarles, T. and D Pederson, “The Spice Page” http://bwrcs.eecs.berkeley.edu/Classes/IcBook/SPICE/

Jason Powell

Utah State University

Langmuir probes, are arguably the most useful and commonly used instrument to diagnose the state of plasma in lab and space experiments. These probes are compact, relatively simple to operate. Over nearly a century, they have been the subject of numerous experimental and theoretical studies to characterize their response to different plasma environments. As a result, there are numerous well theories available to interpret basic plasma state parameters, such as density, temperature, and potential, from currents collected with Langmuir probes. Despite these advances, accurate inferences of plasma parameters from Langmuir probe measurements remain challenging, because of the assumptions made in supporting theories, needed in order to yield simple analytic interpretive formulas. The alternative solution presented here consists of using a combination of computer simulations capable of simultaneously accounting for more physical processes, under more realistic conditions, than what is possible analytically. Results from such simulations are then used to construct synthetic data sets from which inferences techniques can be constructed using adapted multivariate techniques. The approach is illustrated with applications to in situ measurements made on the Swarm, PROBA-2, and NORSAT-1 satellites.

Richard Marchand

University of Alberta

Spacecraft charging results from the transport of electrons and ions from the ambient plasma to spacecraft surfaces. These charged particles carry energy and, depending on the surface material properties, illumination conditions, and electrical connections within the spacecraft, large magnitude potentials can develop. Previous studies have examined whether this differential charging can be exploited to harvest energy from the space environment, assuming unmagnetized spaces plasmas that are stationary with respect to the spacecraft. We present preliminary results for how the areal power density is affected by relaxing these assumptions. Analytic Orbit-Motion-Limited (OML) models and the Spacecraft Plasma Interaction Software (SPIS) are used to determine the energy harvested by a spacecraft in an electron-hydrogen plasma of varying density and temperature.

Sean Young

Stanford University

Missions rely, sometimes exclusively, on analytical verification methods for spacecraft surface charging risk reduction. Tools such as NASCAP2K, SPIS, MUSCAT, EMA3D CHARGE, and others, produce valuable and insightful results when modelling the responses of three dimensional spacecraft in variable and varying space plasma environments. Analysis outputs, notably potentials and electric fields, are strongly dependent on model inputs, including material properties. Over reliance, or naivety, in the use of “default”, incorrectly measured, or supplier datasheet material electrical properties in constructing surface charging simulations can yield inaccurate results. Variability or uncertainty in bulk (dark) conductivity and Secondary Electron Yield (SEY), for example, may obscure subtle influences of the space plasma environment when calculating surface potentials or electric fields. Effects of temperature, radiation dose, contamination, roughness, and non-homogeneity, as examples, have been shown to yield up to several orders of magnitude variability in material electrical properties.

We performed a series of parametric analyses using multiple surface charging tools (NASCAP2K and EMA3D CHARGE) and a simplified 6-sided cube model consisting of a metallic conductor and a single “variably semi-conductive” face. The purpose of these studies was to establish enveloping results for absolute charging (frame charging) and differential charging whilst accounting for uncertainty in inputs. Material property uncertainty included: Orders of magnitude variability in bulk (dark) conductivity and surface resistivity. Correlative variability in Secondary Electron Yield (SEY) properties (delta max and Emax) as shown in Fig. 1; delta max and Emax were varied together to produce credible, physical, SEY curves.Variability / uncertainty in Radiation Induced Conductivity (RIC) was not considered as it does not influence NASCAP2K results.

Also environmental variability: Worst case environments for Single and Double Maxwellians per ISO 19923. Correlative number densities and temperatures in Double Maxwellians as shown in Fig. 2.

Indeed we demonstrated that credible uncertainties or ranges in material properties produce results ranging from acceptable (e.g. little or no differential charging) to unacceptable with hundreds of volts observed. Results to be presented / discussed include: Sources, examples, and ranges of electrical property uncertainty in non-homogeneous and highly insulating materials. Materials such as static dissipative black kapton, Stamet, and carbon composites are ubiquitous in modern spacecraft designs yet their surface charging related electrical properties are not trivial to measure or model. Quantified resultant ranges in absolute (frame) potentials and differential potentials considering the variability in electrical properties (SEY and conductivities). Quantified ranges in absolute (frame) potentials and differential potentials considering the variability in Maxwellian environments. Round-Robin style comparisons between NASCAP2K and EMA3D CHARGE for these basic calculations.

Findings are presented along with initial applications to more complex spacecraft geometries – specifically Van Allen Probes (VAP). These results further illustrate the influence of material electrical properties on spacecraft surface charging. They illuminate the risk in over-reliance on under-informed analytical verification methods and the criticality in properly measured and modelled material properties. These results are of heightened importance when considering, for example, scientific spacecraft with rigorous instrument accommodations requirements, missions with docking and rendezvous operations, missions with electric propulsion systems, and / or carrying human beings.

Justin Likar <justin.likar@jhuapl.edu>

Johns Hopkins University Applied Physics Laboratory

In the laboratory, electrostatic discharges (arcs) on solar array coupons were studied in a simulated GEO environment. Arc current pulses were recorded with a high time resolution - up to 0.2 ns. Spectral properties of these pulses demonstrated no monotonic behavior but rather separated peaks from 100-1000 MHz. Electromagnetic radiation was registered with a Yagi antenna 8 m distant from the arc site and a high speed oscilloscope. Analysis of the spectra shows well-separated “lines”. Electric field strengths were measured near the bell jar in open air. Evaluation of arc current spectra, electric field spectra, and radio frequency spectra revealed statistically significant patterns. Measured radio fluxes only seen in “on-source” scans on GPS satellites and theoretical estimates of the power of electromagnetic radiation from arcs on spacecraft solar arrays, argue in favor of the arcing hypothesis for the source of this radiation. Detrimental effects of arcing in the laboratory were studied with an optical microscope, and multiple burn spots and surface contamination sites were found. These results have enabled a better understanding of the generation of electromagnetic radiation by transient vacuum discharges.

Dale Ferguson

AFRL, Former Senior NRC Fellow

NASCAP is used for creating Spacecraft Surface charging model and analysis for Geo-Spacecraft. NASCAP works in Dos prompt and Inputs are directly taken into DOS-prompt during execution. The input files: Object Definition, Charging Environment, Run Option and Execution command Outputs like Charging model data and time-line data are available in DOS-prompt during execution of the program. The same is saved in text files for offline analysis. After execution of NASCAP, whole data of 20000 sec. is collected from Dos prompt, manually processed in excel and plotted using MATLAB. Due to manual and repetitive steps followed during NASCAP program execution, it requires more time to generate charging model. If any data is missed during execution, whole exercise requires to be repeated.

​Graphical user interface for NASCAP execution using Python, which includes different modules of python viz PySimpleGUI, PyautoGUI and Matplotlib has been developed. GUIDE-NASCAP Interfaces input files e.g., Plasma environment definition, Run option file and Object definition file with the NASCAP software using a GUI. Sunlit-eclipse-sunlit execution are automated with the minimum user input. This software Optimized Execution time for the entire program. Automated execution of “wdraw” and “Termtalk” modules have been incorporated. The pictorial views of spacecraft potentials are obtained using NASCAP module, “OBJPOTL” through a click in GUI. Automated data analysis, image, plots and data table generation with minimum inputs from user are achieved and manual and repetitive steps in conventional method have been eliminated in GUIDE-NASCAP. Processing of data and analysis plots are available immediately. Faster execution and generation of Charging model with minimum user inputs are the prime features of GUIDE-NASCAP.​

This paper details the features of GUIDE NASCAP and validation of Spacecraft charging data using conventional method and Guide Nascap method. The Figures for Automatically generated Object views for analysis, The Absolute Potential Plot generated through Conventional method of NASCAP execution and plotting in MATLAB and the Absolute Potential Plot generated through Guide- NASCAP are attached in Fig-1, 2 and 3 respectively.

Similarly, the Differential Potential Plots have been generated using Conventional method of NASCAP execution, plotting in MATLAB and using Guide NASCAP. The results have been compared and validated.

Anju Damodaran

URRao satellite centre, Indian Space Research Organization