Leonard Kramer

The Boeing Company, Houston, Texas

Thomas W. Kerslake and Joel T. Galofaro

Glenn Research Center, Cleveland, Ohio

Integration Assessment of Visiting Vehicle

Induced Electrical Charging of the International 

Space Station Structure

NASA/TM—2010-216901

December 2010
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Leonard Kramer

The Boeing Company, Houston, Texas

Thomas W. Kerslake and Joel T. Galofaro

Glenn Research Center, Cleveland, Ohio

Integration Assessment of Visiting Vehicle

Induced Electrical Charging of the International 

Space Station Structure

NASA/TM—2010-216901

December 2010

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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NASA/TM—2010-216901 1 

Integration Assessment of Visiting Vehicle Induced Electrical Charging of 

the International Space Station Structure 

Leonard Kramer 

The Boeing Company 

Houston, Texas 77059

Thomas W. Kerslake and Joel T. Galofaro

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The International Space Station (ISS) is a large low earth 

orbit (LEO) satellite. The ISS electrical charging properties in 

this environment affect its operation. The 160 V solar array 

(SA) system on ISS is grounded negative. Small exposed 

conductors on the SA collect electrical charge from the 

environmental plasma driving the metal structure through the 

electrical ground to a negative floating potential (FP) relative 

to plasma. This potential is variable in location and time—a 

result of asymmetric current collection properties moderated 

by local ionospheric conditions. So called v × B induction 

distributes up to 20 additional positive and negative volts 

across ISS depending on its attitude and geographic location. 

Visiting Vehicles (VV) are now a regular part of ISS 

operations. In addition to time-tested Russian Soyuz and 

Progress spacecraft, the European Automated Transfer 

Vehicle (ATV) and the Japanese H-II Transfer Vehicle (HTV) 

have completed missions to ISS. Commercial space transport 

interests, both SpaceX and Orbital Sciences, are moving 

forward with VV development. Both have active plans to dock 

with ISS. The United States NASA VV concept is a 

configuration of the Orion capsule, under development by 

prime contractor Lockheed-Martin Corporation, from the 

Constellation program. 

Most VV, just like the ISS, incorporate SA power systems. 

Any of these VV may experience charging by the same 

process as ISS. Upon physical contact with ISS, the current 

collection properties of a VV combine with ISS and any other 

docked VV. This is a vehicle integration concern as FP must 

be controlled because the vehicle may experience degradation 

of its anodized aluminum surface coatings if electrical hull 

arcing were to occur and an extra vehicular activity (EVA) 

crewman is also exposed to shock hazard. ISS FP is controlled 

by plasma contactor units (PCU) which operate only during 

EVA. 

This report is an assessment of the Orion SA induced 

charging of ISS and motivates a generic limiting electric 

current delivery specification for any docked VV. The Orion 

SA incorporates a negative grounded, 130 V, UltraFlex (ATK 

Space Systems) SA being developed under subcontract by 

Alliant Techsystems Inc. (ATK) of Goleta, California. 

The Orion SA presents small amounts of exposed metal cell 

interconnects and photoactive germanium substrate. To assess 

the current collection characteristic, Orion solar cell test 

coupons were constructed and subjected to plasma chamber 

current collection measurement. Current collected by the 

coupons from the plasma was measured as a bias potential was 

swept between 0 and 120 V. This current voltage sweep data 

determined at the chamber plasma state temperature and 

density (n and T) was scaled to arbitrary geophysical n and T 

according to Langmuir probe theory. The data and the 

theoretical scaling are combined in a numerical model which 

is integrated into the Boeing Plasma Interaction Model (PIM). 

It was found that the SA design for Orion will not affect ISS 

by more than about 2 V during worst case charging conditions. 

A trade study was conducted with PIM to determine that no 

single or combined fleet of docked VV should deliver more 

than 5 mA of current. PCUs can reject up to about 

10 A and therefore can control additional VV if they meet the 

combined docked fleet total of 5 mA. 

1.0 Introduction 

Solar array power systems provide a direct current (DC) 

electrical potential for spacecraft. Designs for such systems 

yield significant efficiencies if power is distributed at high 

voltage. High voltage requires less current and therefore less 

conductor and lower weight. On the other hand, high voltage 

requires greater care in isolating energized elements and 

voltage step down converters are needed for any low voltage 

subsystems. 

On most spacecraft it is customary to ground the negative 

side of the electrical system. Grounding the electrical supply 

polarizes the system providing a means to isolate energized 

conductors to subsystems when power is off. A subsystem 

with an exposed heating element for example is safe 

electrically if the on-off switch breaks the hot side conductor. 

A grounded system also provides a return path for noise 

transients. Nevertheless, it is possible to design systems that 

function with ungrounded power as are used in the standard 

440 V system on United States Navy vessels (Ref. 1). 
NASA/TM—2010-216901 2 

In developing the power distribution system for ISS and 

VV a number of these technical, economic and sometimes 

legal and political requirements are balanced in trade studies 

where the competing requirements are reconciled. A 160 V, 

negatively grounded, DC system was the outcome of a trade 

study performed for ISS. The Orion program, through a 

similar process adopted a 130 V DC design, also grounded 

negative. 

When a VV docks with ISS the respective electrical ground 

of the two vehicles become effectively bonded. At the same 

time the external environment of ISS is an electrically 

conducting plasma and most solar array designs provide 

exposure of electrical conductors to the plasma. Currents will 

generally flow through the plasma and any exposed metal on 

the structure. Negative grounded arrays cause negative charge 

collection on the combined VV-ISS structure contributing to 

accumulated negative electrical potential on the structure. 

Therefore, because the ISS must accommodate mated 

operations with this variety of VV, a technical integration 

assessment of the electrical charging effects from these 

vehicles must be performed. This report describes the special 

challenge involved in developing assessment standards for the 

United States Orion Vehicle. A trade study is also performed 

here which provides general guidelines for the amount of 

current that the ISS can accept safely from any combined fleet 

of docked VV. 

1.1 ISS Electrical Charging Properties 

Solar array electron collection at exposed conductors on the 

positively charged arrays will drive the ISS (or any so 

configured satellite) to a negative potential. Four charging 

scenarios are illustrated in Figure 1. In this figure which has 

been adapted from an unpublished report (Ref. 2) the array 

configuration is abstractly pictured as a single SA oriented 

along the direction that electrical potential is distributed. The 

bonded structure ground is represented by a cube of metal. 

In Figure 1 scenario A, the arrays are depicted in a shunted 

configuration where no potential is supplied. In this scenario 

the structure potential adjusts in response to currents from 

electrons and ions. The electrical potential of the structure 

floats slightly negative. This is the voltage providing 

equilibrium between electrons and ions collected from the 

plasma. 

In the Figure 1 scenario B, the arrays are active. The 

positive voltage distributed across the arrays drives electrons 

through the negative ground until a new equilibrium potential 

is achieved wherein the active collection of electrons at the 

arrays is balanced against collection of ions on the bare metal 

of the structure. 

 

Figure 1.—Illustration of electron and ion collection on solar array and satellite structure driving voltage negative 

relative to external plasma. Shunted arrays (A). Active arrays with no PCU control (B). Active arrays plus Visiting 

Vehicle contribution to electron collection (C). Active arrays with PCU control (D). 
NASA/TM—2010-216901 3 

Our concern regarding VV effect on charging is illustrated 

in Figure 1, scenario C. When mated to ISS the arrays of the 

VV contribute additional exposed conductor at positive 

potential driving the structure to a lower potential than is 

achieved in the unmated scene B. 

In general the lower potential is an operational concern. 

Lower potential applies an electric field across thin (1 to 

15 μm) anodized aluminium on the ISS and can cause it to arc. 

Such arcing is readily observed in the laboratory and the 

characteristic anodized material covers nearly all the nonRussian pressurized and the external truss elements (Refs. 3 

and 4). The anodized coating is carefully designed to facilitate 

optimum thermal protection and such arcing, if not mitigated, 

would degrade and eventually remove the protection. The 

potentials are also a hazard to crew during extravehicular 

activity (Ref. 5). 

1.2 Magnetic Induction (v × B) 

Direct charging of the vehicle from the arrays as described 

in Figure 1 is not the only agent affecting the FP. The motion 

of ISS through the magnetized ionosphere induces an electric 

field drop in the plasma across the linear extent (L) of the 

vehicle structure. It is proportional to the cross product of 

velocity (v) and geomagnetic field vector (B). The induced 

(v  B)·L potential is illustrated in Figure 2. This v  B 

induction moderates electron and ion collection at conducting 

areas of the vehicle. 

The masts supporting the SA are made, in part, of stainless 

steel struts or wires. The wires, although not bonded by 

design, provide metal to metal contact communicating among 

all the struts to the ISS structural metal. The Russian 

pressurized elements also provide an electrically conducting 

mesh on the exposed side of their multilayer insulation 

blankets (Ref. 6). The mesh is intentionally grounded to 

structure and completely shields the Russian elements. 

In darkness electrons collect on the SA masts and on 

exposed conductors on the arrays themselves. The aggregate 

effect of v  B induction on ISS potential results in that slow, 

broad, half orbit structure seen in the FP signature of Figure 3. 

This electron collection tends toward an equilibrium with ions 

recombining primarily at the Russian modules, but also at 

masts and sometimes at the SA. 

1.3 Measurement of ISS Floating Potential 

and Plasma State 

The ISS is equipped with a set of tools developed to 

monitor the plasma state. The instrumentation is the Floating 

Potential Measurement Unit (FPMU) developed for NASA. 

The FPMU is described by Wright et al. and references 

contained therein (Ref. 7). The floating potential (FP) probe, 

among the complement of instruments incorporated in the 

 

Figure 2.—Induced v  B potential in the plasma 

distributed across the structure with PCU not active. 

FPMU, comprises a gold plated spherical shell isolated from 

the bonded structure by an approximately 1011  impedance 

and deployed on a 150 cm boom. 

The electrically conducting nature of the plasma provides a 

well known shielding effect associated with the attraction and 

enhanced concentration of opposing polarity charge carriers in 

the plasma volume near any charge. This shielding property is 

characterized by a length scale known as the Debye length. It 

is effective in isolating the FP probe from disturbance 

associated with nearby charged surfaces. The FPMU 

deployment, in particular, locates the FP probe at least 2 

Debye lengths away from any other structure over the range of 

plasma state conditions achieved on orbit. The FP probe 

sphere therefore floats near the plasma potential providing a 

reference for measurement of the ISS (Ref. 8). 

Figure 3 is an example of FPMU telemetry along with our 

theoretical calculation of FP. The upper panel shows the FP 

probe data and our theoretical calculation for 4 orbits in midNovember of 2009. The abscissa scale is color-coded and 

yellow shows sunlit orbit intervals. We observe that the 

potential so illustrated reflects a negative potential of the 

vehicle structure relative to the FP probe shell. We will 

discuss the theoretical calculation in section 4.0. 

The plasma state comprising electron temperature, T, and 

plasma density, n, are determined by Boeing’s Langmuir 

probe (LP) reduction process. There are two LP; a spherical 

wide bias range and a narrow range cylindrical probe. These 

are known as the Wide LP (WLP) and the Narrow LP (NLP). 

The LP sweep voltage over a range and report the current 

collected from the plasma. The reduction processor determines 

density; n, from the negative bias ion collection region of the 

respective current-voltage (I-V) sweeps. The ion current 

model is then subtracted and temperature is determined from a 

optimal fitting of a appropriate electron current model across 

the near zero bias portion of the I-V sweep. The n and T from 

the WLP and NLP are illustrated in the bottom two panels of 

Figure 3. 
NASA/TM—2010-216901 4 

Figure 3—Example of FPMU telemetry collected November 18, 2009 with corresponding theoretical simulation for

4 orbits of ISS. Red line in the top panel is observed vehicle FP relative to deployed probe. The top panel blue line 

is theoretical FP referred to probe location and based on measured n and T seen in the bottom two panels. 

1.4 Control of ISS Floating Potential 

Early (1990) analysis predicted that the space station would 

float to 130 to 150 V below the plasma potential (Ref. 9). To 

address that projected severe hazard from hull arcing (as well 

as a list of other hazards), plasma contactor units (PCU) were 

developed and deployed on the ISS. These devices expel 

plasma generated from a xenon gas reserve. The effect of the 

PCU is to overcome the voltage gradient supported by a 

plasma sheath and provides a low impedance path from 

structure metal to ionospheric plasma. The PCU redistribute 

negative charge to the environment bringing structure 

potential closer to plasma potential as illustrated in scenario D 

of Figure 1. 

The severe hazard prediction proved erroneous. 

Uncontrolled potentials assumed by the ISS have been much 

less than that needed to initiate arcing contrary to the 

predictions. The state of knowledge at the time of those 

predictions overstated the influence of secondary electron 

emission on solar array collection and seems to have had no 

way to account for all the mitigating effect of ion collection on 

conductors; particularly the Russian mesh mentioned in 

section 1.2. Arcing on the surface of the astronaut’s spacesuit 

or Extravehicular Mobility Unit (EMU), although highly 

unlikely, has been determined to be, and remains, a 

catastrophic hazard. PCUs are therefore operated only during 

EVA to protect the astronaut (Refs. 4 and 5). 

1.5 Visiting Vehicles 

Presently four international interests have or are developing 

spacecraft to dock with ISS. The Russian Soyuz crew vehicle 

and Progress cargo carrier have accumulated decades of 

service including several dozen flights to ISS. The vehicles of 

other nations, notably the European ATV and Japanese HTV 

have docked with ISS and more are planned. Private 

commercial interests are well along in development of 

unmanned vehicles designed to deliver cargo to ISS. 

The private company SpaceX has developed a vehicle 

designated Dragon. The Orbital Sciences vehicle is called 

Cygnus. The Orion vehicle under development by the United 

States is the planned crew transport vehicle following the end 

of Space Shuttle missions in 2011. 
NASA/TM—2010-216901 5 

1.6 Integration of Docked VV to ISS 

All of the space vehicles we discuss here have variable 

design approaches to power generation. Both Russian 

Soyuz/Progress and European ATV vehicles incorporate 28 V 

solar array/battery DC power systems. The Japanese HTV 

power system employs solar arrays that operate at voltages 

near 100 V. Dragon and Cygnus specifications are proprietary 

(although they are disclosed to the ISS program vehicle 

integration authority). 

A formal integration assessment of visiting vehicles by the 

ISS program authority has therefore emerged as a requirement 

for safety and mission success. This report is an attempt to 

advance that agenda. 

2.0 Orion Solar Arrays 

The Orion capsule is outfitted with two disk shaped solar 

array wings seen in Figure 4. The final design study has 

converged on 90 strings of 67 solar cells per wing. The strings 

are laid out on 20 triangular “gore” segments distributed 

among the two wings. One of those gore segments is 

illustrated in Figure 5. 

Figure 4.—The Orion solar array system has 20 “gore” 

segments distributed between two disk shaped wings. 

Figure 5.—Orion solar array “gore” segment highlighting 

one of 90 strings. Working potential of 130 V is 

distributed between the end-terminations and across 

the 67 cells in each string. The grounded end is at 

vehicle potential (VOrion). 

2.1 Characteristics of Orion SA 

Each SA string constitutes a voltaic pile of solar cells. The 

active voltage of approximately 130 V is distributed across the 

pile between the negative grounded point which assumes 

vehicle FP. Although the active side of the cells are covered 

with 3-mil thick cover glass, gaps between cells are available 

to collect current from the plasma as are inter-cell 

interconnects seen in the photograph of a test coupon in 

Figure 6. Additionally, the backside of the cell and its 

electrically conducting substrate are covered with insulating 

material. There are also large metallic “end-termination” and 

“turn-around” connectors shown in Figure 5 which are not 

exposed to plasma. 

2.2 Plasma Chamber Test of Orion SA Cells 

Electric current collection tests on a SA cell coupon were 

conducted as a part of this study. The test coupon, shown in 

Figure 6, was designed and assembled by ATK. It comprises 

16 solar cells arranged in a grid. Testing was conducted in a 

plasma chamber at NASA Glenn Research Center (GRC) 

(Ref. 10). The chamber with the cell coupon installed is seen 

in Figure 7. 

Figure 6.—Solar array chamber test coupon 

highlighting exposed interconnect and covered 

end-terminal and turn-arounds. 

Figure 7.—Solar array coupon in plasma test chamber. 
NASA/TM—2010-216901 6 

Figure 8.—Current collection from Orion solar array 

coupons undertaken by GRC. Dark blue diamond 

(♦) is the baseline design. Red square (□) is from 

special test coupons with exposed metallic end 

terminations. 

The positive and negative terminals were on the test cell 

coupon shorted together and biased externally. The plasma 

state in the chamber was adjusted to a density, n = 

1.01012m–3 and electron temperature, T = 0.2eV to reflect a 

typical ionospheric environment. 

The measurements are shown in Figure 8. The current as a 

function of bias potential, or I-V sweep, from the primary 

baseline design is illustrated as the dark blue diamond (♦)

symbol. 

An alternate test was conducted with the coupon design 

altered to expose metal end turnarounds seen in Figure 6 and 

also exposing the backside of the array. For the alternate test, 

plasma state was adjusted to constitute a higher density; n = 

3.51012m–3 with T = 0.25 eV. The I-V sweep for this variant 

coupon test is shown with the red open square (□) symbol in 

Figure 8. 

The variant exposed metal design is in the nature of a 

control for the purpose of illustrating the effectiveness of the 

end turn-around coverings. We will refer to these as the 

variant exposed metal coupons. 

The current from the variant control is seen to be about two 

orders of magnitude larger than the baseline design. Although 

the density is higher for the test of the variant, most of the 

increased current results from exposed metal on the variant. It 

will be shown to contribute significant charging to ISS. 

3.0 Orion Solar Array Electrical 

Current Model 

The physics governing electron collection of an Orion SA 

and the Langmuir probes are closely related. The Langmuir 

probe is a device used to ascertain plasma state conditions by 

observing the current collected from plasma as the electrical 

potential of the device is varied in time. The ISS Langmuir 

probe measurements provide flight like plasma state (T, n) and 

electric collection characteristics similar to metal surfaces on 

the Orion arrays. In contrast, Orion SA coupon data collected 

Figure 9.—Current collected to Langmuir probe 

deployed on ISS over a wide range of temperatures 

and densities. 

in the GRC chamber are measured at a single plasma state (T0, 

n0). 

Existing plasma Langmuir probe theory is accurate around 

the near zero bias potential (Ref. 11). The theory permits T

and n to be observed but is not directly applicable to high 

positive voltage. The WLP described in section 1.3 provides 

high voltage I-V sweep characteristic curves. On-orbit 

assessment of electric current to the bare metal WLP probe 

biased from –20 to 80 V are available for a wide range of on 

orbit conditions. The data and the theoretical scaling provide 

us with a mathematical scaling: 

I   V;T0 , n0  I V;T, n (1) 

relating the I-V sweep relationship in the chamber; I(V;T0,n0) 

to the sweep at arbitrary T and n. The development of this 

scaling is discussed as follows. 

3.1 Langmuir Probe Current Collection 

Theory 

Figure 9 illustrates the current voltage characteristic for the 

Langmuir probe on ISS. The wide variation in current over 

voltage bias results from the changing T and n as ISS orbits 

earth. The variation can be modeled theoretically in a 

normalization to the electron thermal velocity and the probe 

collection area. 

The normalized current, Ψ and voltage, X are: 

env A

I

ψ

e

Probe  (2a) 

k T

eV X

B

 (2b) 
NASA/TM—2010-216901 7 

Figure 10.—Normalization of the data in Figure 9. 

Where Iprobe is the current collected by the probe and V is the 

bias as in Figure 9. Additional parameters are e; the electronic 

charge (e = 1.6×10–19coulombs), n; the plasma density (m–3). 

A is the probe collection area in m2

, kB is Boltzman’s constant 

(kB = 1.38×10–23 J/K. The term ve is the electron thermal 

velocity e BT/2π e v  k / . A plot of Ψ versus X seen in 

Figure 10 shows that the normalization in Equation (2a) and 

(2b) unifies the variation in the original un-normalized data. 

3.2 On-Orbit ISS Langmuir Probe Data 

The red line in Figure 10 is a model of the current given by: 

  0 7

ψ X 0.6 1 X .    (3) 

With the definition of Ψ we obtain: 

  0.7 IT n,T  e  n  veA  1 X (4)

which then provides a basis for rationalizing the transform 

expressed in Equation (1). We use Equation (4) as a proxy for 

the actual on-orbit response to temperature and density. 

3.3 Model of Orion Current Collection 

The basis for the scaling in Equation (1) is contained in the 

proportional variation resulting from arbitrary density, n and 

electron temperature, T, as in: 

     

   T 0 0 0 0

T

M I n ,T ,X 0 Ψ V,n ,T

I n,T,X 0 Ψ V,n,T I V,n,T I V      (5) 

Here, IM is the measured current taken from the data in 

Figure 8. It is implemented as a lookup table functionally 

dependent on bias potential (IM = IM(V)). The term in the 

numerator represents the model current variation at any 

arbitrary V, n and T whereas the denominator provides the 

model at the chamber state. A rearrangement provides the 

Orion electron collection model: 

  

0 7

0

0 0

M

T

V 1

T

V 1

T

T

n

n I V, n,T I V

.





























   (6) 

where the temperature T is expressed in eV. We see in 

Equation (6) that density scales current in direct proportion to 

n. Temperature variation is more complicated; related to 

empirical power law fit obtained in Equation (3). Equation (6) 

is the Orion current model. 

4.0 ISS Plasma Interaction Model 

In this section we present a brief description of the Plasma 

Interaction Model (PIM) developed by Boeing. 

We introduce PIM here as a tool to evaluate the floating 

potential of the ISS (Ref. 12). The algorithm expresses 

accumulation of total electric charge from models of the 

electron and ion currents on conducting surfaces and from 

results of a space charge calculation at the active solar arrays 

performed by the Science Applications International 

Corporation (SAIC). The important contribution from SAIC is 

a finite element solver solution for a potential barrier 

generated by high voltage (0 to 160 V) at the solar arrays 

(Ref. 13). 

PIM is a non-steady state solver based on a principle of 

electric charge conservation. The model incorporates 

subsystem model currents from three components; the ISS SA, 

the exposed conductors on the Russian element and from the 

SA masts. 

Although the Russian elements appear to be covered by 

multilayer insulation blankets which ordinarily exhibit 

dielectric properties, the outer, exposed layer of these blankets 

is embedded with a silver coated copper mesh which is 

bonded to the structure ground (Ref. 6). 

The subsystem models, including the Russian blanket 

model are parametrically controlled by the vehicle potential 

and plasma state (n and T) as well as attitude and orientation 

of the vehicle. The algorithm uses plasma state properties of 

density, n, and temperature, T, determined from on orbit 

FPMU Langmuir probe instrumentation. The lower two panels 

in Figure 3 show the plasma n and T from the Boeing 

Langmuir probe reduction processor. The upper panel 

illustrates measured and theoretical FP. The theoretical FP is 

referred, through the (v  B)·L calculation described in 

section 1.2, to the FP probe location. The comparison between 

the FP probe measurement and the PIM model output 

validates the PIM algorithm for the purpose of this study. 

The Orion model of Equation (6) is configured in the PIM 

software along with the existing models of the solar array, 
NASA/TM—2010-216901 8 

Russian blankets and array masts. The FP for a characteristic 

variation in plasma state during a 1 hr time period is simulated 

in PIM. The period includes eclipse exit and is shown in 

Figure 11. Four different scenarios are assessed and we will 

discuss each in turn. 

4.1 Baseline (No Docked VV) Floating 

Potential 

In the No Orion, scenario A, represented by a green solid line 

in Figure 11, the PIM floating potential is evaluated as a 

baseline for comparison with the active model Orion currents. 

For this baseline and the three model scenarios to be discussed, 

the plasma state (n, T) seen in the lower two panels of Figure 11 

are taken from a well known International Reference 

Ionosphere. The No Orion case is a typical variation of on-orbit 

conditions as seen in the measurements of Figure 3. 

The charging of the Orion spacecraft is of greatest concern 

at eclipse exit where the spacecraft emerges from darkness and 

the arrays are in full ram; oriented toward the sun. As the orbit 

evolves, the sun orientation separates from the ram direction 

and electron collection is mitigated. The ISS SA also 

undergoes array regulation as batteries charge and system 

power demands vary. 

The current from the three components in PIM are 

illustrated in Figure 12. It is seen that in darkness the solar 

array and the conducting elements on the Russian segments 

contribute a net positive current balanced approximately by 

electron collection on the stainless steel wires of the SA masts. 

At eclipse exit, the SA become active and immediately 

dominate electron collection pushing the masts together with 

the Russian segments to the positive current side of the 

collection model. Our primary concern therefore is the 

charging at eclipse exit where all ISS SA are contributing to 

electron collection. 

4.2 Integration of Orion Model into PIM 

To evaluate charging properties of Orion the model 

characterized by Equation (6) was incorporated as one more 

subsystem in PIM using the GRC I-V sweep data from 

Figure 7 to characterize IM(V) in Equation (6). 

For the duration and location in the ISS orbit we adjust the 

ISS FP to account for Orion current, adding up the current 

contribution through each cell while simultaneously adjusting 

for its individual contribution to the working potential and its 

local v  B induction. We accumulate this sum all the way up 

to the positive terminal of the string. 

Active Orion SA scenarios illustrated in Figure 11 include a 

single Orion designated case B and a Dual Dock case C. We 

will discuss first however the exposed metal case D 

characterising the electron collection properties of the Orion 

SA using plasma chamber data collected and seen in Figure 7 

from test coupons prepared with exposed metal end 

terminations. 

Figure 11.—PIM modelled floating potential for 4 

docking configurations. Plasma state from the IRI 

model is illustrated in the bottom two panels. 

Figure 12.—PIM modelled currents contributing to ISS 

charging with No Orion; the case (A) in Figure 11. 

The severe charging approaching –80 V seen in Figure 11 

for the exposed metal variant SA case indicates that the end 

termination covering for the baseline design is effective in 

protecting ISS from charging. Currents modelled from the 

exposed metal coupon are illustrated in Figure 13. There is an 

immediate transient at eclipse exit. During the transient 

current balance is not maintained; corresponding to a rapid 

accumulation of several hundred milli-coulombs if negative 

charge on the ISS. 
NASA/TM—2010-216901 9 

Figure 13.—PIM modelled currents using the 

exposed metal control data set. 

Figure 14.—PIM modelled currents with a baseline 

single docked Orion. 

The currents, amounting to on the order of 50 mA, supplied 

by the exposed metal in this simulation overwhelm all other 

sources of negative current. This explains the rapid jump in 

floating potential seen in the simulation illustrated in 

Figure 11 for the exposed metal SA case. We emphasize that 

at eclipse exit current balance is not maintained and the metal 

on the vehicle accumulates a net negative electric charge of 

about 500 milli-coulombs until equilibrium in current balance 

is re-established after a few seconds. 

Figure 14 on the other hand illustrates Orion currents for 

the single docked baseline design adopted for the integrated 

vehicle. We observe that the modelled currents from the Orion 

are near zero which explains the close conformance of FP 

model result seen in Figure 11 between the baseline and single 

dock Orion result. 

4.3 Simulation Result—Negligible Orion SA 

Impact 

The Orion model results seen in Figure 11 illustrate that the 

Orion array design produces negligible impact on ISS 

charging for this typical scenario. An audit of the result 

indicates that the Dual Dock Orion (trace C in Figure 11) 

produces less than 2 V extra charging. 

Figure 15.—Summary of trade study showing that 

5 mA limit contributes negligible impact to mated 

ISS - VV configuration. 

5.0 Trade Study 

The simulation result described in section 4.0 contains 

uncertainty because the Equation (6) model is derived from 

Langmuir probe characteristics having near optimal current 

collection characteristics. In contrast, the arbitrary orientation 

and exposure of metal and semi-conductor on the Orion SA 

coupons presents an uncertain and complex electron collection 

characteristic. While the simulation, based on plasma chamber 

measured current suggests that the Orion design is acceptable, 

it is necessary for us to provide an objective standard to assess 

VV generically. To this end we have undertaken a trade study. 

In the trade study we set aside the complex VV modeling 

strategy of Equation (6) that uses the data from Figure 7 and 

invoke instead an analysis to establish the VV induced current 

limit that yields negligible impact to ISS-VV integrated 

vehicle. 

For the trade study the Orion model was modified to deliver 

a simple constant current when illuminated in sunlight and 

zero current when in eclipse. The parameter was varied among 

5, 10, and 50 mA. This variation is consistent with the order of 

magnitudes seen in the simulation results for the exposed 

metal case of Figure 13 and the single docked baseline Orion 

case of Figure 14. The trade study result is shown in 

Figure 15. It is seen that when current delivered by a VV is 

5 mA or less negligible additional charging is experienced by 

ISS. 

6.0 Summary and Conclusions 

The electrical power system design for the ISS and the 

various VV address requirements imposed by obligations 

associated with the particular vehicle architecture. As we have 

pointed out, when VV dock with ISS the electrical grounds are 

joined and the current collection characteristics contribute to 

the aggregate electrical floating potential of the combined ISS 

and VV. 

 
NASA/TM—2010-216901 10 

We have studied theoretically the electrical charging 

contribution from the baseline Orion solar array and also from 

a special variant with exposed metal end turnarounds. The 

result of the calculation indicates that the baseline Orion 

arrays are a well integrated design and will contribute no more 

than about 2 V under worst case conditions to ISS electrical 

charge. The simulation from the exposed metal variant 

produced unacceptable charging of ISS approaching –80 V. 

A trade study was undertaken to provide an objective 

criterion for VV integration into ISS. It was found that a 

5 mA limit on current provided by VV made less than 2 V 

contribution to ISS floating potential under worst case 

conditions and it is proposed to be an acceptable electric 

current for design purposes. 

6.1 Orion Design Characteristics Contributing 

to Electrical Charging Relief 

The covering of the metal turnarounds seen in Figure 6 and 

covering of the photoactive substrate on the back of the cells is 

found to provide relief of ISS charging concerns. There 

remain metallic intercell interconnects and some exposed 

substrate in gaps between cells but these are not contributing 

significant current under the most adverse conditions 

encountered. 

6.2 Multiple Docked VV Concerns 

The ISS provides docking ports for multiple VV and we are 

proposing that 5 mA is a design criterion to limit adverse 

charging effects. This raises the challenge of apportioning 

5 mA to any combined fleet of docked VV. The resolution of 

this matter is a subject of continuing study. 

7.0 References 

1. United States Navy, Interface Standard for Shipboard 

Systems Section 300A Electric Power, Alternating 

Current, Military Standard (MIL-STD) 1399, section 

300A. 

2. Katz, I., Gardner, B., Mandell, M. “ISS Plasma Arc 

Dynamics,” Report to ISS PCU Team 1 Technical 

Interchange Meeting, June 21, 2000, NASA/GRC, 

Cleveland, Ohio. 

3. Carruth Jr., M.R., Schneider, T., McCollum, M., 

Finckenor, M., Suggs, R., Ferguson, D., Katz, I., 

Mikatarian, R. Alred, J., Pankop, C., “ISS and Space 

Environment Interactions in Event of Plasma Contactor 

Failure,” Spacecraft Charging Technology, Proceedings, 

Noordwijk, Netherlands, 95-100, 2001. 

4. Kramer, L., Leung, P.L., Schneider, T.A., Vaughn, J.A., 

Tyler R. Black, T.R., Hamilton, D.R., Steve Koontz, 

“Internal EMU Resistance Impact on Suit Arcing During 

EVA,” 44th AIAA Aerospace Sciences Meeting and 

Exhibit 9–12, Reno, Nevada, AIAA–2006–870, 2006. 

5. Kramer, L., Hamilton, D., Mikatarian, R., Thomas, J., and 

Koontz, S., “Positive Voltage Hazard to EMU Crewman 

from Currents through Plasma,” Fourth IAASS 

Conference, Huntsville, Alabama, May 19-21, article in 

press, ESA Special Publication SP-680, 2010. 

6. Schneider, T., Vaughn, J., Minow, J., Hwang, K.-S., 

Wright, K., Reddell, B., Alred, J., Mikatarian, R., Leung, 

P., Koontz, S. “Current Collection Characteristics of 

International Space Station Russian Thermal Blanket 

Material,” AIAA–2006–867, 44th AIAA Aerospace 

Sciences Meeting and Exhibit, Reno, Nevada, Jan. 9–12, 

2006. 

7. Wright Jr., K.H., C.M. Swenson, D.C. Thompson, A. 

Barjatya, S.L. Koontz, T.A. Schneider, J.A. Vaughn, J.I. 

Minow, P.D. Craven, V.N. Coffey, L.N. Parker, T.H. Bui, 

Charging of the International Space Station as Observed 

by the Floating Potential Measurement Unit: Initial 

Results, IEEE Trans. Plasma Sci., 36(5), 2008. 

8. Swenson, C.M., D. Thompson, C. Fish, The Floating 

Potential Measurement Unit, AIAA–2003–1081, 41st 

Aerospace Sciences Meeting and Exhibit, 6–9 January, 

2003. 

9. Ferguson, D.C., Snyder, D.B., Carruth, R., Findings of the 

Joint Workshop on Impacts of Space Station Freedom 

Ground Configurations, NASA Tech. Memo. 103717, 

1990. 

10. J.T. Galofaro, B.V. Vayner, and G.B. Hillard, 

“Experimental Charging Behavior of Orion UltraFlex 

Array Designs,” Journal of Spacecraft and Rockets, 

Vol. 47, Number 3, May-June 2010, pp. 521–532. 

11. Brace, L.H., “Langmuir Probe Measurements in the 

Ionosphere,” Measurement Techniques in Space Plasmas: 

Particles Geophysical Monograph 102, 23–35, 1998. 

12. Mikatarian, R., H. Barsamian, J. Alred, J. Kern, J. Minow, 

S. Koontz, Electrical Charging of the International Space 

Station, 41st Aerospace Sciences Meeting and Exhibit, 6–

9 January, AIAA–2003–1079, 2003. 

13. Mandell, M.J., Davis, V.A., Jongeward, G.A., Gardner, 

B.M., “ISS Solar Array Electron Collection Model 

Report,” Report submitted to Boeing Company by 

Science Applications International Corporation, May 23, 

2002. 
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Integration Assessment of Visiting Vehicle Induced Electrical Charging of the International 

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14. ABSTRACT

The International Space Station (ISS) undergoes electrical charging in low Earth orbit (LEO) due to positively biased, exposed conductors on solar arrays that collect electrical 

charges from the space plasma. Exposed solar array conductors predominately collect negatively charged electrons and thus drive the metal ISS structure electrical ground to a 

negative floating potential (FP) relative to plasma. This FP is variable in location and time as a result of local ionospheric conditions. ISS motion through Earth’s magnetic field 

creates an addition inductive voltage up to 20 positive and negative volts across ISS structure depending on its attitude and location in orbit. ISS Visiting Vehicles (VVs), such 

as the planned Orion crew exploration vehicle, contribute to the ISS plasma charging processes. Upon physical contact with ISS, the current collection properties of VVs 

combine with ISS. This is an ISS integration concern as FP must be controlled to minimize arcing of ISS surfaces and ensure proper management of extra vehicular activity 

crewman shock hazards. This report is an assessment of ISS induced charging from docked Orion vehicles employing negatively grounded, 130 volt class, UltraFlex (ATK 

Space Systems) solar arrays. To assess plasma electron current collection characteristics, Orion solar cell test coupons were constructed and subjected to plasma chamber 

current collection measurements. During these tests, coupon solar cells were biased between 0 and 120 V while immersed in a simulated LEO plasma. Tests were performed 

using several different simulated LEO plasma densities and temperatures. These data and associated theoretical scaling of plasma properties, were combined in a numerical 

model which was integrated into the Boeing Plasma Interaction Model. It was found that the solar array design for Orion will not affect the ISS FP by more than about 2 V 

during worst case charging conditions. This assessment also motivated a trade study to determine acceptable plasma electron current levels that can be collected by a single or 

combined fleet of ISS-docked VVs.

15. SUBJECT TERMS

Plasma interactions; International Space Station (ISS); Crew exploration vehicle; Solar arrays 

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