48th International Conference on Environmental Systems ICES-2018-181 8-12 July 2018, Albuquerque, New Mexico

International Space Station Charging Environments: Modeling, Measurement, and Implications for Future Human Space Flight Programs

Steve Koontz1, Emily Willis2, John Alred3, Erica Worthy4 NASA Johnson Space Center, Houston, Texas, 77058 & NASA Marshall Space flight Center, Huntsville, Alabama, 35811

William Hartman 5, Benjamin Gingras5, William Schmidl5 The Boeing Company, Houston, TX 77059 USA

The International Space Station (ISS) orbital altitude and inclination (~400 km and 51.6o) determines the natural factors affecting ISS spacecraft charging: 1) high speed flight through the geomagnetic field, 2) electrical power system operations and interactions with the relatively cold, high-density ionospheric plasma and, 3) exposure to energetic auroral electrons at high latitude. In this paper we present the results of 12 years of ISS spacecraft charging measurements, using the ISS floating potential measurement unit (FPMU), and compare the measurement results with numerical modeling of ISS charging processes. The ISS spacecraft charging environment is radically different from those encountered at higher altitudes in ’s magnetosphere and in cis-lunar space, where energetic charged particles (space radiation) are the more important natural factors affecting spacecraft charging. We evaluate the applicability of ISS spacecraft charging management methods and experience to future human beyond LEO.

Nomenclature C/NOFS = Communication/Navigation Outage Forecast System eV = Electron Volts FP = Floating Potential FPMU = Floating Potential Measurement Unit GEO = Geosynchronous ISR = Incoherent Scatter Radar LEO = Ne = Ionospheric electron density in electron number per cubic meter Ni = Ionospheric ion density in ion number per cubic meter NLP = Narrow (sweep) Langmuir Probe PV = Photovoltaic PIP = Plasma Impedance Probe Te = Electron (plasma) temperature UT = Universal Time (Greenwich Mean Time) UV = Ultraviolet radiation WLP = Wide (sweep) Langmuir Probe

I. Introduction pacecraft charging hazards are caused by the accumulation of electrical charge on spacecraft and spacecraft S components produced by interactions with space plasmas, energetic charged particle populations, and solar UV photons as well as operation of spacecraft electrical power systems, propulsion systems, and payloads.1-3 Spacecraft

1 ISS System Manager for Space Environments, NASA/JSC 2101 E NASA Pkwy, Mail Code ES4, Houston, TX,77058 2 Space Environments Team Lead, NASA/MSFC, Redstone Arsenal, Mail Code EV44, Huntsville Al, 35812. 3 Deputy Chief, Material sand Process Branch, NASA/JSC 2101 E NASA Pkwy, Mail Code ES4, Houston, TX,77058 4 ISS Deputy System Manager for Space Environments, NASA/JSC 2101 E NASA Pkwy, Mail Code ES4, Houston, TX,77058 5 Engineer/Scientist, Boeing ISS Space Environments, 13100 Space Center Blvd., Mail Code HB3-20, Houston, TX 77059

International Conference on Environmental Systems charging hazard effects include avionics systems soft (recoverable) anomalies as well as hard failures of spacecraft avionics and electrical power systems and have led to the partial or complete loss of some spacecraft.1-3 Electrical charge can accumulate on: 1) the exterior surfaces of the spacecraft (surface charging), 2) in dielectrics interior to the spacecraft (internal or, 3) deep dielectric charging, as well as in the spacecraft conducting structure (frame, structure, or absolute charging).4-7 Differential charge accumulation in or on different parts of the spacecraft can produce local voltage gradients that lead to dielectric breakdown arcs with accompanying current and voltage pulses of sufficient magnitude to interfere with spacecraft operations or damage spacecraft materials and electronics. The location and magnitude of charge accumulation in or on the spacecraft depends on the spacecraft mechanical and electrical configuration (especially spacecraft or component capacitance), spacecraft electrical systems operation, spacecraft materials electrical properties, and how the spacecraft and it’s components interact with the natural spacecraft charging environments consisting of the local space plasma, energetic charged particle populations, and solar illumination (i.e., photoelectron emission) environments.4-7 Free electrons and UV photons are the natural spacecraft charging environment factors that often dominate spacecraft charging processes. The kinetic energy of the free electron population expressed in electron volts (eV) is one of the most important factors determining spacecraft charging voltages in some charging environments, e.g. (GEO), Earth’s geo-tail, the polar 800 km Earth orbit, and in planetary radiation belts. Spacecraft charging voltages on the order of 1 kV to >10 kV have been observed in these environments as have numerous spacecraft charging anomalies and failures.8, 9 Spacecraft charging is modeled quantitatively using variations of the charge balance equation which sets the sum of positive and negative charging currents, to the object under study, to zero at steady state.4-7 The International Space Station (ISS) orbital altitude and inclination (~400 km and 51.6o) determine the ISS spacecraft charging environment, which is radically different in many respects from those encountered at higher altitudes in Earth’s magnetosphere (GEO, geo-tail, and the radiation belts), in cis-lunar space, and on the lunar surface. The population of high kinetic energy electrons (KeV to MeV) characteristic of the higher altitude regions of the magnetosphere and cis-lunar space, and responsible for high voltage spacecraft charging in those regions, are largely absent from the ISS low-earth orbit (LEO) environment, except in the auroral zone at high latitude 10, 11. The kinetic energy of the free electron populations at higher altitudes in earth’s magnetosphere is an important factor determining spacecraft charging and charging mitigation. ISS flies in the F2 region of earth’s ionosphere, a relatively cold (average electron kinetic energy ~ 0.1 eV), high- density (1010 to 1012 e-/m3), plasma, that can mitigate surface charging on forward facing surfaces. The Floating Potential (FP, the electrical potential difference between an object and the surrounding plasma) of a small metallic object in the ISS orbit is only a few volts negative because electrons are much easier to collect than ions, on account of the much lower electron mass, and the net current collected by the object goes to zero at a small negative voltage10, 11. Spacecraft charging in LEO is often driven by voltages and currents generated by the spacecraft itself. In the case of ISS, high speed flight of the large ISS conducting structure through the geomagnetic field can lead to voltage gradients of up to 0.50 V/m along the 100 m ISS truss, at high latitudes, produced by the Lorentz Force (aka motional electromagnetic force or EMF). Motional EMF is calculated using a variant of the Lorentz equation applicable to metallic objects moving relative to a stationary magnetic field, i.e., V = v x B . L, where V is the voltage (FP) difference between the two ends of the conducting structure of length L, v is the ISS velocity vector, B is the geomagnetic field vector, L is the length of ISS conducting structure, and x is the vector cross product operator10, 11. Exposed metallic surfaces grounded to ISS conducting structure collect ionospheric current proportional to the magnitude of the ISS motional EMF, which depends on ISS location in the geomagnetic field The photovoltaic (PV) power system generates photovoltaic string voltages of ~ 160 V when in sunlight. The role of the 160 V PV power system in ISS spacecraft charging has been previously described10, 11. Briefly, electron and ion collection from the ionosphere by the powered PV strings balances when the FP of the string is predominantly negative 10, 11. Both the photovoltaic cell interconnects in all PV cell strings and the current collection busses on each PV array wing are encapsulated in dielectric, so that the exposed metallic PV circuit elements that usually make the largest contribution to spacecraft charging in LEO are unable to collect current from the ionospheric plasma, effectively mitigating much of the PV array driven charging for ISS. However, ISS PV array current collection is still a significant contributor to ISS charging as a result of current collection on PV cell edges exposed under the glass PV cell coverslips. The internally generated ISS voltages, motional EMF and PV system string voltage, drive collection of ions and electrons from the ionosphere producing a predominantly negative ISS conducting structure FP10, 11. Typically, ISS flies in what is referred to as the +XVV flight attitude with the velocity vector parallel to the pressurized element axis and perpendicular to the 100 meter truss (Figure 1), leading to higher motional EMF voltages

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International Conference on Environmental Systems at high latitudes where v x B is maximized. The contribution of photovoltaic string voltage to ISS charging is primarily dependent on the orientation of the active surfaces of the PV arrays with respect to the ISS velocity vector and the solar vector. The ISS PV arrays can track the during each orbit on two axes. PV array wing current collection from the ionosphere and contribution to ISS FP, depends on the PV wing being in sunlight to generate voltage, and facing forward (i.e. facing down the velocity vector) to collect electrons and ions10, 11. Plasma current collection by ISS PV array wings exhibits an important angle of attack effect so that current collection drops off significantly as the angle between the velocity vector and PV array active surface normal approaches 90 degrees 10, 11. For angles of attack greater than 90 degrees, current collection by the PV aray effectively drops to zero as a result of the well-known LEO spacecraft wake effects 12, 13.

Figure 1. ISS in the +XVV flight attitude. The velocity vector (v) is parallel to the pressurized element stack and perpendicular to the 100 meter long (L=100 m) truss as shown to the left. The motional EMF interaction of ISS with the geomagnetic field at high northern latitudes is shown to the right.

In this paper we present the results of 12 years of ISS spacecraft charging measurements, using the ISS floating potential measurement unit (FPMU)15-17, and compare the measurement results with numerical modeling of ISS charging processes. ISS is a large metallic structure and flight through the geomagnetic field at orbital speed dominates ISS charging. Collection of ionospheric electrons by the large 160 V PV arrays is the next largest contributor. Charging by auroral electrons is detectable but makes a relatively minor contribution. We also report the observation of short duration (~1 sec) Rapid Charging Events (RCE) often associated with shunt/un-shunt operations of the 160 V PV arrays, a phenomena not predicted before flight. Finally, we present an assessment of the extent to which ISS spacecraft charging observatons and management experience can be of value to future programs in cis-lunar space and beyond where spacecraft charging environments and effects are often radically different than those that dominate ISS charging in LEO15-17.

II. The Floating Potential Measurement Unit (FPMU)

A. Instrument description and measurement principles The ISS carries the Floating Potential Measurement Unit (FPMU), a suite of four instruments that measure the electron number density (Ne), electron temperature (Te) and the potential of the ISS structure relative to the plasma environment at the FPMU location (i.e. the Floating Potential (FP))15 . These instruments include: 1) Wide-sweep Langmuir Probe (WLP), 2) Narrow-sweep Langmuir Probe (NLP), 3) Floating Potential Probe (FPP), 4) Plasma Impedance Probe (PIP). For over 10 years (CY 2006 to present), the FPMU has been collecting ISS charging data. It

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International Conference on Environmental Systems is installed on Camera Port 6 on the P1 truss segment, but was previously installed on the S1 truss segment until November 21, 2009. Each plasma characteristic is measured by at least two instruments for redundancy and validation. Both the NLP and the WLP provide Ne (electron density) and Te (electron temperature) data. The NLP is a cylindrical probe that uses the floating potential measured by the FPP as input to perform a +/- 5V sweep around the floating potential. This sweep allows for a high resolution measurement of the Langmuir Probe I-V curve in the electron and ion retarding regions, which is then used to calculate the plasma density and temperature. The WLP is a spherical probe that provides the same type of measurement as the NLP, but with a wider sweep from -20V to +80V and lower voltage step resolution. Both the WLP and NLP measure the Ne and Te at 1 Hz. Both measure the Te in the 0.05 to 0.4 eV range, while the Ne is measured in a range of 109/m3 to 5 x 1012/m3. However, the PIP measures the Ne at 512 Hz and has an effective range of 108/m3 to 1013/m3. The FP is measured by the WLP at 1 Hz and the FPP at 128 Hz18. The FPMU does not measure ion species concentrations, electric fields, magnetic fields, or drift speeds. The Current FPMU location on ISS is shown in Figure 2. A drawing of the FPMU with dimensions and probe configurations is shown in figure 3. After deployment and activation of the FPMU on ISS during August of 2006, the measurement accuracy and precision was confirmed and validated by comparison with independent ionospheric measurements made with the ground based incoherent scatter radar at the Massachusetts Institude of , Haystack Observatory in Westford, Massachusetts19, the Communication/Navigation Outage Forecast System (C/NOFS) satellite20, and the University of Massachusetts at Lowell, Ionosphere Real-Time Assimilative Model (IRTAM)21, that uses input data from a global network of Digisonds (ionospheric vertical radiofrequency sounders) to produce a global map of ionsopehric density at altitude data every 15 minutes.

Figure 2: The current Location of FPMU on ISS. Camera port 6 on the P1 truss segment

Figure 3: The Floating Potential Measurement Unit. FPMU dimensions and probe locations. 4

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B. In-flight instrument validation – Millstone Hill Incoherent Scatter Radar (ISR) On UTC day 220 of the year 2006, ISS flew over the Millstone Hill incoherent scatter radar site while the radar was making vertical sounding measurements of ionospheric temperature and density. FPMU measurements of ionospheric Ne and Te made at UTC 8.7 hours and UTC 10.8 hours are in reasonably good agreement with ISR measurements made during the two overflights as shown in figure 3. FPMU ionospheric density and measurements are plotted against UTC time and the Millstone Hill measurements during the overflights are marked by the circled X symbols.

Figure 3: FPMU measurements of ionospheric density and temperature compared to Millstone ISR measurements made during an ISS overflight. Red refers to ionospheric temperature and blue to ionospheric density.

C. In-flight instrument validation – Spacecraft Conjunctions Electron density (Ne) data taken from both the NLP and WLP has been compared to Ni measurements from the Coupled Ion-Neutral Dynamics Investigation (CINDI) on the Communication/Navigation Outage Forecast System (C/NOFS) satellite. Earth’s ionosphere is electrically neutral, on average, so that Ni ad Ne should be nearly equal (ref). We found 151 conjunctions where the ISS and the C/NOFS orbital paths crossed the same latitude and longitude within a 50 km in altitude and 15 minute in time. Of those conjunctions, 45 occurred when both the FPMU and CINDI were collecting Ne/Ni data. The data for both the FPMU NLP and the FPMU WLP showed good agreement with the CINDI data as shown in figure 4.

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Figurre 4: C/NOFS Coupled Ion-Neutral Dynamics Investigation (CINDI) Instrument Ni measurments compared with FPMU Langmuir Probe (WLP andNLP) Ne measurements.

D. In flight instrument validation - Ionosphere Real-Time Assimilative Model (IRTAM) We investigated alternative methodologies for obtaining ionospheric data in the event of an FPMU failure. The University of Massachusetts at Lowell analysis tool Ionosphere Real-Time Assimilative Model (IRTAM) was identified to provide the data. IRTAM provides the data through their Global Assimilative Modeling of Bottomside Ionospheric Timelines (GAMBIT) database. A detailed description of IRTAM is given by Ref. 21,and a brief summary is given here. The International reference Ionosphere (IRI) model of Earth’s ionopshere requires certain details in order to calculate Ne at a given location. These features include the altitude and Ne for the E-peak, F1-peak, and F2-peak, along with the bottomside parameters B0 (thickness), B1 (shape) and D1 (F1 thickness). Using this information, a set of diurnal/spherical coefficients is determined that are used for the calculations (http://giro.uml.edu/GAMBIT/GAMBIT_CoefficientsMessageFormat.pdf). In the IRI model, these parameters are climatological because they are based on long term average historical data. Therefore, they do not determine the present environment. IRTAM adjusts these coefficients every 15 minutes in order to produce a best fit to real-time bottomside (below the F2-peak) measurement data from the Global Ionospheric Radio Observatory (GIRO): a worldwide network of digital radio frequency vertical ionospheric sounders (Digisondes). We used these real–time coefficients to calculate the Ne along the ISS orbital track. All FPMU data from CY 2006-2017 has been compared to IRTAM calculations. As expected, the IRTAM correlates better than simply using IRI since it uses real-time input data (see figure). Both the IRI and the IRTAM have a maximum probability that is observed to be slightly biased negatively in comparison to FPMU data. The digisonde data used as input for the IRTAM is concentrated mainly over the continents, with very few locations on islands to provide information about the ionosphere above the oceans. We investigated if this inhomogeneous distribution affected the IRTAM accuracy by only comparing IRTAM results with FPMU measurements taken over the continents. IRTAM predictions over the land are significantly more accurate. In this case 80% of the IRTAM Ne values are within +/-30% of the Ne measured by the FPMU.

E. Validity of single point measurement for generation of ISS voltage maps. In the latitude range the ISS orbits, the geomagnetic field lines both connect back to the Earth. Generally this is a closed system. In this region neutral winds can drive ions both perpendicular and parallel to the magnetic field depending on the altitude, while polarization electric fields produced by dynamo processes affect the plasma motion perpendicular to B so as maximize v x B in the +XVV flight attitude. These electric fields are in the 1 mv/m range, so they can be ignored when calculating the potential at different locations on the ISS. However, the ISS motion through the geomagnetic field produces a significant electric field in the ISS reference frame at mid and high latitudes (~0.5 v/m). These electric fields can cause the plasma potential to vary by 40 V from the starboard side to the port side of the truss. The FP at any location on the ISS can be easily calculated by using the FP at the FPMU, the cross product

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International Conference on Environmental Systems of the ISS velocity and the geomagnetic field, and the position vector between the FPMU and the new location because ISS grounding and bonding requirements make the spacecraft essentially one conducting structure.

Figure 5: Comparison of IRTAM and IRI Ionospheric Density Predictions with FPMU Measurement

III. ISS Spacecraft Charging Measurements Figures 6 and 7 show examples of ISS FP measured by the FPMU with the FP features often observed. ISS was in the +XVV flight attitude for these measurements. As ISS enters sunlight at the eclipse exit point of its orbit, the solar arrays are charging ISS batteries and are completely unregulated. As ISS batteries approach a fully charged state, solar array downregulation comes into play and blocks of photovoltaic cell strings are removed from service by shunting (shorting out the PV strings) so as not to overcharge the batteries. At eclipse exit, and before downregulation comes into play, the solar arrays are in an optimal state and configuration to collect ionospheric electrons causing a more negative floating potential. Solar array eclipse exit driven charging can be seen at UT hour 2.2 and 3.8 in figure 6 and at UT hour 20.7 in figure 7. The gradual rise and drop between 2.1 UT and 2.8 UT also between 2.8 UT and 3.5 UT) in figure 6 and between 20.1 UT and 20.4 UT as well as between 21.2 UT and 21.9 UT in figure 7 are the result motional EMF driven current collection by exposed conducting surfaces on ISS, driven by motional EMF voltage in ISS conducting structrue. Peaks in the FP are sometimes observed as the ISS flies through high Ne regions like the Appleton Anomaly (figure 7 at 21.1 UT) if the PV arrays are still facing forward and illuminated at orbital noon, and there are often small drops in the FP as the solar arrays are no longer illuminated and facing wake as the ISS enters eclipse. Motional EMF: This is partially covered in the section describing the validity of the FPMU measurements for calculating the FP at any location on the ISS. As shown in figure 1, the motional EMF will cause the FP to vary as the ISS flies through the magnetic field. However, it also causes the FP to be dependent on the location in the ISS frame (see figure 8). In fact, when the FP is rising on the Starboard side of the ISS, the EMF can be causing the FP on the port side to drop. Figure 8 shows a voltage map of the ISS at a high latitude. This ISS FP map was calculated using the FP measured by the FPMU, the ISS velocity and flight attitude, and the geomagnetic field vector.

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Figure 6: ISS Charging over an Orbit with PV arrays facing forward at eclipse exit

Figure 7: ISS Charging over an Orbit with ISS PV arays in a different configuration

PV array electron collection in sunlight: The solar arrays: Each SA contains 82 strings, and each string has 400 cells. The potential across a solar array string (V) is currently set at 160 V while in insolation. Unlike other current collecting areas on the ISS which are directly exposed to the plasma, the solar array current-collection is done in small gaps between the cover glass. The cells are 8x8 cm, whereas the gaps that run along the sides of the cells are only 32 mil (0.08 cm) wide. The potential in the gap is controlled by the ISS FP and the bias of the solar cells. However, the FP on the cover glass surface is controlled by the ion and electron currents collecting on the surfaces coming to equilibrium. Due to the thermal velocity of the electrons being much higher than the ram velocity of the ISS, the electron current for the glass surface is omni directional. However, the thermal velocity of the ions is much lower than the ram velocity of the ISS because of their mass. This causes the resulting coverglass potential when the ion and electron currents equilibrate to have a large dependency of the solar arrays angle to ram. When the solar arrays are

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International Conference on Environmental Systems facing ram, the ions a more able to remove charge that the electrons have deposited on the glass due to the maximum flux. Therefore, the FP of the glass is low and current can more easily collect on the conducting surfaces in the gaps. When the solar arrays are edge on, ions cannot as easily remove charge deposited by the thermal electrons, resulting in a stronger field that limits the amount of current that can reach the conducting materials in the gaps.

Figure 8: ISS VxB . L Floating Potential Map at high latitude with the velocity vector parallel with the pressurized element stack

Dependence on ionospheric density and temperature: Higher Ne produces a higher ion current capable of removing the charge that shields the gaps. Therefore, higher Ne results in higher charging. However, high Te results in higher electron velocity and charges the glass more, producing a more affective field to block the current from reaching the gaps and limits the charging. FPMU data has shown that ISS charging at eclipse exit is correlated with Ne and anticorrelated with Te. Rapid Charging Events: In addition to the standard floating potential profiles caused by motional EMF and array electron collection in sunlight, transient charging events have been observed that are believed to be related to the charging of the ISS dielectric surfaces. The first transient events were observed in 2006 and were called Rapid Charging Events (RCE), which occur during local plasma depletions22-28. These events are larger in magnitude, extreme cases approaching negative 70 Volts, and faster in duration than the normal eclipse exit charging. They are characterized by rise times on the order of seconds and fall times of tens of seconds. They occur at eclipse exit when the arrays enter sunlight and correlate with lower plasma densities. An example of an RCE compared to a normal eclipse entry event is shown in Figure 8. Both of these events occurred on the same day, August 13th, 2016, but the RCE occurred during a period of lower density. The density during the RCE was very low at approximately 7x109 m- 3, while the density during the normal profile was 5x1010 m-3. Starting in 2010, new transients were observed which that are larger and faster than the RCE events. These events occur when the solar arrays are un-shunted in full sunlight, which is not a standard mode of operation but has occurred during off-nominal operations and anomalies. One example of un-shunt transient event data taken during these operations is shown in Fig. 9 and a closeup of the first un-shut event is shown in Figure 10. The arrays were unshunted in full sunlight as part of a planned operation on July 27th, 2010. A floating potential transient occurred for each of the 8 arrays. These transient events can approach negative 100 Volts. They have rise times of milliseconds followed by a fall on a timescale of tens of milliseconds22-28.

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International Conference on Environmental Systems Plasma responds to any changes in the electrical environment with both resistive and capacitive properties27. Any time an object with a different potential than the plasma is submerged in the medium a transition region for the potential (i.e. the sheath) is formed in order to control the current an object can collect. The current through the outer sheath boundary must equal the current collected on the objects surface. If the current crossing the sheath through random thermal motion does not match what is collected on the object, then the sheath expands until the surface area of the sheath is large enough that the 2 currents balance27. This expansion is an example of Maxwells displacement current, and is what causes the quick rise in the floating potential. The relaxation of the floating potential occurs when the current becomes capable of flowing through an alternate path (either inductance no longer has a large effect, or the resistivity of the medium has dropped), and the sheath relaxes to its normal state. Higher density plasmas have lower resistivites and, therefore have shorter relaxation times. An example of the observed dependence of RCE relaxation times on iosospheric plasma density is shown in figure 11.

Figure 8: Normal eclipse exit and rapid charging compared

Figure 9:Transient Array Unshunt Events on July 27th, 2010

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Figure 10: Close up of Transient Array Un-shunt Charging Event on July 27th, 2010

Figure 11: Dependence of rapid charging event relaxation time on ionospheric density

Low ionospheric density and high latitude (auroral) charging events: Spacecraft charging events, caused by precipitating auroral electron streams, have been commonly observed on in the 800 km altitude polar earth observation orbit29. Auroral charging events have been observed in the FPMU data during eclipse at high latitudes. These events typically occur during geomagnetic storm conditions when the auroral oval extends down to the maximum ISS latitude. One example is shown in Figure 11. This charging event occurred during G2 level geomagnetic storm conditions as set by the Space Weather Prediction Center [https://www.swpc.noaa.gov/]. At the time of the charging event, the ISS was at its highest latitude and in darkness. The increase in net negative current collection due to the high energy precipitating auroral electrons causes detectable, but small, negative floating potential transients.

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Figure 12: Transient event recorded as ISS transits the southern auroral zone on September 3rd, 2012.

The auroral charging event that occurred between UT hours 7.94 and 8.02 of the 86th day of 2008 (2008/86) produced the largest negative auroral charging FP observed on ISS to date. The 2008/86 auroral charging event data are shown in figure 12. The ISS was in the auroral zone during for 144 seconds during the 2008/86 event; however the times when the FP was rising (i.e., when ISS experienced discrete auroral charging events) were much shorter (~12 seconds).These events also correlate with local electron density (Ne) enhancements caused by collisional ionization of the plasma as shown in the WLP and NLP Ne plots in figure 12.

Figure 13: FPMU FP (dark blue) and WLP (light blue) and NLP (red) electron densities

Defense Meteorological Satellite Program (DMSP) data (GMT 2008_86) show a large frequency of current densities above 2x10-5 A/m2 along the ISS 2008/86 charging event flight path (http://www.ospo.noaa.gov/Operations/DMSP/) as shown in figure 13. The red line in figure 13 (corresponding to 144 seconds of flight time) displays the ISS trajectory during the auroral charging event where current densities can exceed 2x10-5 A/m2.

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International Conference on Environmental Systems The assumption in the new model of current collection on ISS anodized Al materials (auroral electrons can penetrate the 30 micron chromic anodize coatings on the anodized aluminum MM/OD shields on ISS US/JAXA and ESA pressurized elements) to produce conducting structure charging and is supported by the timelines and magnitudes of current densities observed by DMSP on that date. The extremely thin anodic coating of the relatively large ISS pressuriszed elements is the reason that ISS capacitance is so large, and the large capacitance is one of the reasons that ISS charges to an FP only about -18 V in environments where smaller spacecraft with much smaller capicitances, would be expected to charge to FP values on the order of -1000 V. A simple “toy model” calculation can help clarify the situation. By definition, V = Q/C, where V is voltage, Q is total capacitor charge and C is the capacitance. Q depends on the current density, the current collecting area, and the exposure time. Assuming an auroral charging current of 2 x 10-5 amps/(m2 sec) for a duration of 10 second and estimating electron collection areas equal to the cross sectional area for both the ISS pressurized elements (C = 1.1 x 1010pF) and a 1 meter diameter spherical satellite with 10 micron dielectric film (C = 1.26 x 106 pF) it is straightforward to show that an FP of -2000 V is expected for the smaller spacecraft but only -13 V for ISS, a value reasonably close to the observed -18 V FP observed during the ISS auroral charging event o 2008/86.

Figure 14: The red line (corresponding to 144 seconds of flight time) displays the ISS trajectory where current densities can exceed 2x10-5 A/m2.

IV. Modeling and Prediction of ISS Spacecraft Charging Phenomena: The Plasma Interaction Model (PIM) The Plasma Interaction Model (PIM) is a time dependent charging model that calculates the ISS FP using the ISS capacitance (0.011 F) and currents from multiple sources15-17. The total ISS current collection is calculated by taking the sum of 3 main current sources: 1) Mast Wires (MW), 2) Solar Arrays (SA), 3) Russian Segment (RS), and 2 optional current sources, 4) the Plasma Contactor Units (PCU), and, if the vehicle is berthed/docked, 5) the Japanese space agency HTV unmanned cargo vehicle. The MW, SA, and HTV in the PIM software tool are capable of collecting both ions and electrons. However, the RS is always an ion collector because it is located at the center of the truss where the FP is always negative and has no biasing (i.e. solar arrays). The PCUs only expel a negatively charged plasma while in discharge. Once the total current is determined for a time interval, the new FP is calculated by the approach represented in the following equation 1 V  I (V )dt V ISS  ISS ISS ISS0 CISS

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International Conference on Environmental Systems Here IISS is the total current collected on the ISS, CISS is the ISS capacitance, VISS0 is the initial ISS FP at the PIM software tool reference point, and VISS is the new FP at the PIM software tool reference point. Notice that the variable IISS is a function of VISS which changes with time. This dependency requires a unique approach to solving for VISS that will not be discussed in this paper. Each current collecting area is at a different location, hence, at a different time dependent FP with respect to the plasma. The PIM software tool uses the International Geomagnetic Reference Field (IGRF) model along with the ISS attitude in order to calculate the geomagnetic field (B) in the ISS reference frame. That B is then used to determine the potential (V) applied in the current calculations at each point by a simple translation.

V = VISS + (v x B) ∙ L

PV array driven charging was discussed above. In the PIM software tool, the gaps are treated as imaginary surfaces at a particular potential lying between two cover glass surfaces that are held at another potential determined by the Ne and Te and solar array angle to ram. The superposition of those potentials determines the potential field produced above the gaps which is then used to calculate the current collection. ISS FP values calculated using the PIM charging model and FPMU measurements of ionospheric Ne and Te are compared with FPMU FP measurements in figures 14, 15, and 16. Figures 14 through 16 show scatter plots of FPMU measurements and PIM model predictions of ISS FP. The measurements were taken between dates 2007/188 and 2013/105. A total of 2328 FPMU data points are represented. Figure 14 represents data points as recorded at the FPMU location on ISS while figures 15 and 16 represent the FP values expected port and starboard truss tips. If PIM model calculations were in perfect agreement with FPMU measurements then all the points would fall along the 45 degree regression line. Inspection of figures 14 -16 shows that the majority of the data points are within a few volts of the regression line there are a number of outliers with FPMU FP values much larger than the FP values calculated using PIM. Closer examination of the data set revealed that all the outliers were examples of the “rapid increase in potential” or RCE peaks discussed in Section III. While some rapid charging peaks in negative FP have caused in the natural environment, many if not most are the result of electrical power systems operations or anomalies. Photovoltaic power system anomalies (power-on-reset anomalies) and commanded full solar array wing shunt/un-shunt commands can both produce transient charging peaks of up to -100 V amplitudes. These events are actively being researched and preliminary models have been developed which show they are sensitive to plasma temperature, density, array ram angle, dielectric charging characteristics, and solar array turn-on time 22-29.

Figure 15: ISS FP at the FPMU location and FP values calculated using the PIM ISS charging model

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Figure 16: ISS FP at the port Truss tip location and FP values calculated for that location using the PIM ISS charging model. Nominal charging data points are dark blue diamonds and rapid charging peaks are magenta squares

Figure 17: ISS FP at the Starboard Truss tip location and FP values calculated for that location using the PIM ISS charging model. Nominal charging data points are dark blue diamonds and rapid charging peaks are magenta squares

V. Applicability of ISS Spacecraft Charging Experience to Future Human Exploration Beyond LEO

The specific combination of ionospheric plasma and geomagnetic field environments that dominate the ISS natural spacecraft charging environments occur nowhere else in the inner . Earth and are the only in the inner solar system with significant planetary magnetic fields and Mercury’s magnetic field strength is only about 1% of Earth’s30. Relatively cold high density ionospheres only exist around and in the inner solar system and approach the density of Erath’s ionosphere only at relatively low altitudes; below 420 km for Venus and below 15

International Conference on Environmental Systems 200 km for Mars31. Aurorae have, however, been observed on Mars and Venus, despite the absence of global magnetic fields, and related magneto tail electron acceleration and precipitation processes should be expected on Mercury also, though the spacecraft charging threat, if any, has not been characterized for any of these planets32, 33. With the exception of occasional ISS exposure to auroral electron charging events, natural spacecraft charging environments at higher altitudes in Earth’s magnetosphere and in cis-Lunar space are radically different from those encountered by ISS in LEO. The natural spacecraft charging environments in cis-lunar space is dominated by energetic electrons and solar UV photons and the GEO spacecraft charging environment has been selected by NASA for design and verification of spacecraft destined for operations in cis-lunar space. The cross program natural space environments for design document SLS-SPEC-159 has been baselined by System (SLS), Orion, and Gateway and specifies a worst case GEO spacecraft charging environment for and verification even though the programs using the specification spend very little time in GEO or the geo-tail itself. Spacecraft in cis-lunar space will be operating in the relatively benign solar wind spacecraft charging environment most of the time, however, GEO like charging environments are expected in Earth’s geo-tail, during solar particle events and the passage of solar coronal mass ejections. It should be noted that when the is full as viewed from Earth the Moon is in Earth’s geo-tail and geosynchronous charging conditions have been observed in that environment by multiple spacecraft including Lunar Prospector, Themis, and Artimus34-41. NASA and Department of Defence standards and guidelines are radically different for the LEO and GEO charging environment3. GEO spacecraft charging mitigation relies heavily on the selection of materials with appropriate electrical properties along with electromagnetic compatability and electromagnetic interference (EMI/EMC) design and testing standards that support mitigation electrostatic charge build-up and electrostatic discharge risks3. Active, on-demand, surface charging control has also been utilized on spacecraft in GEO and related charging environments3. Materials and methods that have been used successfully in the ISS LEO environment are not recommended for use in cis-lunar or interplanetary environments3 without critical evaluation and testing to assure safety, reliability, and mission success. However., ISS operations in LEO do provide important lessons learned supporting risk management and reduction for future human spaceflight operations beyond LEO. Because high-voltage large-area PV arrays are baselined at present for the cis- platform and the Mars Transport that will depart from Gateway, large floating potential excursions are expected to result from power-on-resets and/or commanded PA array full wing shunt/un- shunt commands and any risks to gateway or Mars transport presented by the expected rapid charging peaks must be identified and mitigated during the spacecraft design and verification process. Observations of the ionospehric density dependence and array configuration on the magnitude and duration of ISS FP excursions on full PV wing shunt/un- shunt operations suggest that operational controls of Gateway or Mars Transport could be used to mitigate the risk presented by the subject shunt/un-shunt operations. For example, operation of the electric propulsion, thereby creating an artificial ionosphere around the vehicle, may mitigate the risk. However, more research is needed to fully understand how these high voltage solar array operations affect the FP and spacecraft operation in these environments.

VI. Conclusion The ISS orbit presents a set of natural spacecraft charging environments that, with very few and very specific exceptions, are not likely to be encountered in near future human exploration programs beyond LEO. Spacecraft FP voltages induced by ISS flight through the geomagnetic field and operation of the 160 V PV arrays cause collection of currents from an otherwise benign ionospheric environment leading to charging of ISS conductive structure. There are no naturally occurring charging environments containing a comparable ionosphere or geomagnetic field in the inner solar system. Spacecraft in the cis-lunar space environments will spend most of their operating time in a relatively benign solar wind environment but will be exposed to GEO like charging conditions intermittently. Because the Moon is in the geo-tail when the moon is full as viewed from earth, GEO like charging conditions and the risk of geomagnetic storm effects can be expected for about 6 days out of every lunar month. GEO like spacecraft charging conditions can also be expected during passage of coronal mass ejections and solar particle events. However ISS does provide important experience and lessons-learned with respect to spacecraft induced charging environments. ISS observations of large FP voltage transients associated with full PV wing shunt/un-shunt and power- on-reset anomalies point to significant possible risks for any solar electric spacecraft in cis-lunar space using high- voltage, large-area photovoltaic power systems. ISS experience with the effects of ionospheric density and array operation and configuration on the duration of the transient voltage events also points to possible operational mitigation of the risks and continuing research is needed in this area.

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