ISO 11221:2011

INTERNATIONAL ISO
STANDARD 11221
First edition
2011-08-01
Space systems — Space solar panels —
Spacecraft charging induced electrostatic
discharge test methods
Systèmes spatiaux — Panneaux solaires spatiaux — Matériaux d'essai
de décharge électrostatique induite par la charge du vaisseau spatial

Reference number
©
ISO 2011
©  ISO 2011
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ii © ISO 2011 – All rights reserved

Contents Page
Foreword . iv
1  Scope . 1
2  Terms and definitions . 1
3  Symbols and abbreviated terms . 5
3.1  Symbols . 5
3.2  Abbreviated terms . 7
4  Tailoring . 7
5  Test Items . 7
6  Preliminary tests for ESD inception statistics . 10
6.1  Purpose . 10
6.2  Test facility . 10
6.3  Test coupon . 10
6.4  External circuit . 10
6.5  Test procedures . 11
6.6  Estimation of number of ESD events in orbit . 11
7  Qualification test for secondary arc . 12
7.1  Purpose . 12
7.2  Triggering method and test facility . 12
7.3  External circuit . 12
7.4  CIC gap test — Test coupon and procedures . 13
7.5  Panel test — Test coupon and procedures . 13
7.6  Success criteria . 13
8  Characterization tests for robustness to ESD and plasma interaction . 14
8.1  Power degradation . 14
8.2  Secondary arc . 15
8.3  Power leakage to plasma . 15
8.4  Solar array back surface test . 17
9  Test report . 17
Annex A (informative) Plasma interaction and electrostatic discharge effects on solar array . 19
Annex B (informative) Secondary arc qualification processes . 22
Annex C (normative) Chamber size for a test using LEO-like plasma . 23
Annex D (informative) ESD events analysis . 24
Annex E (informative) Spacecraft charging analysis . 27
Annex F (informative) Derivation of theoretical surface flashover current . 29
Annex G (normative) External circuit of secondary arc test . 31
Annex H (informative) Solar cell I-V characteristics measurement . 36
Annex I (informative) Secondary arc statistics . 38
Annex J (informative) Solar array back surface test . 41
Bibliography . 42

Foreword
ISO (the International Organization for Standardization) is a worldwide federation of national standards bodies
(ISO member bodies). The work of preparing International Standards is normally carried out through ISO
technical committees. Each member body interested in a subject for which a technical committee has been
established has the right to be represented on that committee. International organizations, governmental and
non-governmental, in liaison with ISO, also take part in the work. ISO collaborates closely with the
International Electrotechnical Commission (IEC) on all matters of electrotechnical standardization.
International Standards are drafted in accordance with the rules given in the ISO/IEC Directives, Part 2.
The main task of technical committees is to prepare International Standards. Draft International Standards
adopted by the technical committees are circulated to the member bodies for voting. Publication as an
International Standard requires approval by at least 75 % of the member bodies casting a vote.
Attention is drawn to the possibility that some of the elements of this document may be the subject of patent
rights. ISO shall not be held responsible for identifying any or all such patent rights.
ISO 11221 was prepared by Technical Committee ISO/TC 20, Aircraft and space vehicles, Subcommittee
SC 14, Space systems and operations.
iv © ISO 2011 – All rights reserved

INTERNATIONAL STANDARD ISO 11221:2011(E)

Space systems — Space solar panels — Spacecraft charging
induced electrostatic discharge test methods
1 Scope
This International Standard specifies qualification and characterization test methods to simulate plasma
interactions and electrostatic discharges on solar array panels in space. This International Standard is
applicable to solar array panels made of crystalline silicon, gallium arsenide (GaAs) or multi-junction solar
cells. This International Standard addresses only surface discharges on solar panels.
2 Terms and definitions
For the purposes of this document, the following terms and definitions apply.
2.1
active gap
gap between solar cells across which a potential difference is present when the solar array power is available
2.2
blow-off
emission of negative charges into space due to an electrostatic discharge
2.3
collisionless plasma
plasma in which the mean free paths of electron-neutral, ion-neutral and coulomb collisions are longer than
the scale length of interest
NOTE Chamber length is an example of a scale length of interest.
2.4
differential charging
spacecraft charging where any two points are charged to different potentials
2.5
differential capacitance
capacitance between any two points in a spacecraft, especially between the insulator surface and the
spacecraft body
2.6
differential voltage
potential difference between any two points in a spacecraft during spacecraft charging, especially between the
insulator exterior surface potential and the spacecraft chassis potential
2.7
discharge inception voltage
lowest voltage at which discharges of specified magnitude will recur when a DC voltage is applied between
any two points in a spacecraft, especially between the insulator surface and the spacecraft body
2.8
electrical breakdown
failure of the insulation properties of a dielectric, resulting in a sudden release of charge with possible damage
to the dielectric concerned
2.9
electric propulsion
spacecraft propulsion system in which the thrust is generated by accelerating charged particles that are
neutralized before they are ejected in order to produce a jet
2.10
electrostatic discharge
electrical breakdown of dielectric or gas or vacuum gaps, and also of surface interface of dissimilar materials,
caused by differential charging of parts of dielectric materials and their interfaces
2.11
gap distance
distance between biased cells or conductors
2.12
glow discharge
gaseous discharge with a surface glow near the cathode surface
NOTE The origin of the ionized gas is mostly ambient neutral gas molecules rather than metal vapour from the
cathode surface.
2.13
inverted potential gradient
inverted voltage gradient
result of differential charging where the insulating surface or dielectric reaches a positive potential with respect
to the neighbouring conducting surface or metal
NOTE This phenomenon is also known as PDNM (positive dielectric negative metal).
2.14
non-sustained arc
passage of current from an external source through a conductive path that lasts only while the primary
discharge current flows
See Figure 1.
2.15
normal potential gradient
normal voltage gradient
result of differential charging where the insulating surface or dielectric reaches a negative potential with
respect to the neighbouring conducting surface or metal
NOTE This phenomenon is also known as NDPM (negative dielectric positive metal).
2.16
permanent sustained arc
passage of current from an external source through a conductive path that keeps flowing until the external
source is intentionally shut down
See Figure 1.
NOTE Some permanent sustained arcs may leave a permanent conductive path even after the shut-down.
2.17
Poisson process
stochastic process in which events occur independently of one another
2 © ISO 2011 – All rights reserved

2.18
power generation voltage
potential difference between the positive and negative terminals of a solar array string
2.19
primary arc
trigger arc
developed phase of a primary discharge, under an inverted potential gradient, which is associated with
cathodic spot formation at a metallic or semiconductor surface
2.20
primary discharge
initial electrostatic discharge which, by creating a conductive path, can trigger a secondary arc
See Figure 1.
NOTE The current can include blow-off current and surface flashover current.
2.21
punch-through
dielectric breakdown between two sides of an insulator material
2.22
ram
space in front of and adjacent to a spacecraft in which the plasma density can be enhanced by the motion of
the spacecraft
2.23
satellite capacitance
absolute capacitance
capacitance between a satellite body and the ambient plasma
2.24
secondary arc
passage of current from an external source, such as a solar array, through a conductive path initially
generated by a primary discharge
NOTE Figure 1 shows the various stages of a secondary arc.
2.25
snapover
phenomenon caused by secondary electron emission that can lead to electron collection on insulating
surfaces in an electric field
2.26
solar array front surface
solar array surface where solar cells are laid down
NOTE Solar cells are laid down on the side of a solar panel that normally faces the sun.
2.27
solar array back surface
solar array surface where solar cells are not laid down
NOTE Solar cells are not laid down on the side of a solar panel that normally faces away from the sun.
2.28
surface charging
deposition of electrical charges onto, or their removal from, external surfaces
2.29
surface flashover
surface discharge propagating laterally over a dielectric material
NOTE Surface flashover is sometimes called a “brushfire discharge”.
2.30
temporary sustained arc
passage of current from an external source through a conductive path that lasts longer than a primary
discharge current pulse but terminates without leaving a permanent conductive path
See Figure 1.
2.31
wake
trail of rarefied plasma left behind by a moving spacecraft

Key
1 primary discharge (blow-off + flashover)
2 non-sustained arc (NSA)
3 temporary sustained arc (TSA)
4 permanent sustained arc (PSA)
5 secondary arc
i current
I short-circuit current of one or more solar array circuits
sc
t time
The primary discharge is fed by absolute and differential capacitances. The secondary arc is fed by the solar
array power.
Figure 1 — Stages of secondary arc
4 © ISO 2011 – All rights reserved

3 Symbols and abbreviated terms
3.1 Symbols
A area of surface of plasma
s
C bypass capacitance
BC
C differential capacitance
CG
C capacitor representing capacitance between solar panel structure and ambient plasma
GS
C capacitor representing capacitance underneath the cells through the Kapton layer
kapton
C satellite capacitance
sat
C capacitor representing capacitance of solar array string
string
C capacitance per unit area of coverglass
v
C capacitor representing capacitance of solar array string and capacitance underneath the cells through
the Kapton layer
C capacitor representing capacitance of solar array string and capacitance underneath the cells through
the Kapton layer
C capacitor representing capacitance of solar array string and capacitance underneath the cells through
the Kapton layer
D fast switching diode
D fast switching diode
D fast switching diode
d sheath thickness
sh
I reverse saturation-current density, in amperes per square metre (A/m )
I power supply representing power generated by the solar array
I short-circuit current of one or more solar array circuits
sc
I current of a solar array section
section
I current of a solar array string
string
i current
j number of bins
k Boltzmann constant
L inductance to form the pulse current shape
ext
n diode constant
P probability that an event occurs in the i-th bin
i
Q charge
q elementary charge
R resistance
R resistance to form the pulse current shape
ext
R resistance to adjust the voltage between two strings under test
L
R U /I resistance needed to get the right voltage and current in the loop simulating the solar
section section section
panels section
R U /I resistance needed to get the right voltage and current across the solar cells simulating
string string string
the solar array string under arcing test
r radius of plasma
T temperature, in kelvins (K)
T
e electron temperature
T ion temperature
i
t time
t time to threshold differential voltage
ESD
U constant current source
U constant voltage source
U voltage of a solar array section
section
U voltage of a solar array string
string
V power supply representing charging potential of spacecraft body
b
v velocity of plasma wavefront
p
V potential difference
 angle
 Debye length
D
ρ electron density
e
ρ ion density
i
 coverglass potential
CG
 satellite body potential
sat
6 © ISO 2011 – All rights reserved

3.2 Abbreviated terms
19
eV electron volt (1 eV = 1,602  10 J)
CIC coverglass interconnect cell
ESD electrostatic discharge
GEO geosynchronous orbit
IPG inverted potential gradient
LEO low Earth orbit
NPG normal potential gradient
NSA non-sustained arc
PA primary arc
PEO polar Earth orbit
PI plasma interaction
PSA permanent sustained arc
TSA temporary sustained arc
4 Tailoring
Specifications described in this International Standard are tailorable upon agreement between the customer
and the supplier.
5 Test items
NOTE Annex A provides an overview of the subject of spacecraft charging and electrostatic discharge (ESD)
phenomena for readers who are not familiar with the subject.
The aims of the plasma interaction (PI) and ESD tests are to simulate the detrimental phenomena anticipated
in space for a given solar array design, to evaluate a design’s resistance to the phenomena and to provide
data necessary for the judgment of qualification and characterization.
Figures 2 and 3 present the test items specified in this International Standard, with flow charts to summarize
the logic flow of each test. The purpose of a preliminary test for ESD statistics is to define the statistics helpful
for selecting the test parameters (such as the number of primary discharges inflicted upon a test coupon),
defining the margins of the test parameters and defining the confidence level of the test results. If proper
statistics for these numbers and probabilities are already available, the preliminary test is not required for the
qualification of secondary arcs. Annex B provides a brief rationale of the structure of the flow chart in Figure 2.
Figure 2 — Logic flow of ESD tests
8 © ISO 2011 – All rights reserved

Figure 3 — Logic flow for determining the necessity of a power leakage characterization test
6 Preliminary tests for ESD inception statistics
6.1 Purpose
The purpose of this test is to characterize the ESD (primary discharge) inception threshold in terms of
differential voltage between the coverglasses and the solar array circuit. This differential voltage can be used
as a tool to estimate the number of ESD events during the mission lifetime in orbit.
6.2 Test facility
The test facility shall be able to simulate the charging processes of a solar array insulator in orbit. If the solar
array is for a GEO satellite, the solar array insulator shall be charged using either an energetic electron beam
3
or UV irradiation, or a combination of both, in a vacuum chamber with a pressure lower than 1  10 Pa
6
(7,5  10 Torr). The electron energy shall be less than 30 keV so that the charging takes place mostly over
the insulator surface, and not below it. The vacuum chamber for a geosynchronous orbit (GEO) solar array
test shall be equipped with an adequate device to determine the insulator charging potential, such as a
non-contacting surface potential probe, preferably mounted on an (x)-(y) scanning device.
If the solar array is for a low Earth orbit (LEO) spacecraft, the solar array insulator shall be charged by a low-
energy plasma with a temperature below 10 eV in a vacuum chamber with a pressure that guarantees a
collisionless plasma. If the solar array is for a polar Earth orbit (PEO) spacecraft and auroral electrons are
responsible for differential charging, the solar array insulator should be charged using an energetic electron
beam. If the solar array is for a PEO spacecraft and low-energy ionospheric ions are responsible for
differential charging, the solar array insulator should be charged using a low-energy plasma. See Annex C for
the minimum chamber size for a low-energy plasma test.
The test facility shall be equipped with a device to record an adequate image of the test coupon during the test
so that ESD locations can be identified either during or after the test.
6.3 Test coupon
The test coupon(s) shall consist of at least three strings of three cells to represent a cell surrounded by other
cells. The test coupon(s) should
a) reflect the production variation with respect to parameters that can affect the ESD inception threshold,
such as degree of grouting, coverglass overhang, cell spacing, etc. on the total number of cells on the
test coupon(s),
b) include all the features of a flight panel, such as bus bars, through-holes, terminal strips, wire harness,
hold-down, etc.,
c) include the mitigation techniques that represent the flight model as closely as possible, if the solar panel
design involves ESD mitigation techniques such as a dissipative coating, and
d) consider the worst condition during the life of the spacecraft, such as after thermal cycling, repaired cells,
and other conditions that can lead to a greater risk of ESD and secondary arcs.
6.4 External circuit
In the test, the vacuum chamber serves as the circuit ground. If the charging situation in space is the inverted
potential gradient, bias the test coupon to a negative potential with a DC power supply. If the charging
situation is the normal potential gradient, ground the test coupon. (See Figure 4 for a circuit diagram.) A small
amount of capacitance may be connected to the DC power supply if a brighter flash of ESD light is needed to
identify its location. Limit the capacitance so that the electrostatic energy dissipated does not cause
degradation of the solar cells on the test coupon(s). An energy of less than 5 mJ is recommended. As the
capacitance of a coupon alone sometimes exceeds the limit, external capacitance should not be used for a
large coupon of more than about 20 cells. To record the ESD in this event, use a sensitive camera.
10 © ISO 2011 – All rights reserved

a) Inverted potential gradient b) Normal potential gradient
Key
1 coupon
2 vacuum chamber
Figure 4 — Test set-up for the ESD inception test
6.5 Test procedures
If an electron beam gun or a UV source is used for charging the test coupon, the test shall be carried out until
a statistically significant number of ESD events, no less than 10, occur on the test coupon. The test coupon
surface potential shall be measured repeatedly during the test by a non-contacting surface potential probe.
The coupon surface potential closest to each ESD location shall be identified and recorded. The minimum
differential voltage is the minimum value among all the ESD events recorded. Be aware of the uncertainties
associated with the spatial resolution of the probe and the temporal variation of the potential depending on the
time of measurement from the ESD inception.
If a low-energy plasma source is used, the differential charging voltage can be approximated by the chamber
plasma potential, which is usually positive by several times the electron temperature, minus the negative
coupon potential. The uncertainty is in the order of the electron temperature. The coupon bias voltage shall be
varied to cover all the possible charging potentials in orbit. In the case of PEO spacecraft, the waiting time at
each bias voltage should be no less than 20 min. In the case of LEO spacecraft, the waiting time at each bias
voltage should be no less than 90 min. At low bias voltages where the probability of ESD is very low, a longer
waiting time is recommended to improve the statistics. See Reference [1] for an example of characterizing the
arc rate per unit time under a low-energy plasma environment. If the threshold is unknown, plot the arc rates
at different bias voltages on a logarithmic scale and find the voltage where the probability of an arc over a
given time becomes negligible, assuming that ESD inception is modelled as a Poisson process (see
Reference [2] for an example).
6.6 Estimation of number of ESD events in orbit
It can be useful to analyse the number of ESD events expected in orbit as a basis for discussion to determine
the number of primary discharges in the subsequent tests. See Annex D for details. Other methods of analysis
may also be used to compute the number of ESD events.
7 Qualification test for secondary arc
7.1 Purpose
The purpose of this test is to qualify a given design of solar panel for flight. The purpose of the coverglass
interconnect cell (CIC) gap test (7.4) is to demonstrate that no damaging secondary arc occurs even when
primary discharges are forced to occur directly on the active CIC gap, which is a possible worst case scenario.
The purpose of the panel test (7.5) is to show that no damaging secondary arc occurs, even after a significant
number of primary discharges occur all over the flight-representative test coupon. See Annex B for a more
detailed rationale of both tests.
It may not always be possible to replicate the orbital worst case scenario in the laboratory experiment (e.g. low
temperature at the end of eclipse, the effects of thermal cycling on gap distance, failure or aging of grouting,
and the outgassing time period in orbit). However, every effort shall be made to simulate the worst case
scenario or to extrapolate the test results to represent the worst case scenario parametrization.
7.2 Triggering method and test facility
If it can be confirmed that the probability of a transition from a primary discharge to a secondary arc does not
depend upon the method of primary discharge inception, any method may be used to cause primary
discharges, irrespective of the anticipated charging situation in orbit. If the transition probability depends upon
the testing plasma environment, the same test facility as used for the primary discharge inception threshold
test shall be used. In either case, the shape and amplitude of the primary discharge current in the test shall be
plausibly representative of the current expected in orbit.
There is a risk of primary discharge and subsequent secondary arcs in LEO even for a GEO spacecraft as it
passes through LEO during the orbit transfer. Also, if a plasma emission device is used, such as a plasma
contactor, LEO-type arcing may occur when the device is first turned on or off. Therefore, for a GEO
spacecraft, the test should be performed under the conditions in a LEO-type plasma in addition to the GEO-
type test. See Annex C for the minimum chamber size for a low-energy plasma test.
The test shall take place under vacuum in a test chamber with a pressure that guarantees the physical state of
3
a collisionless plasma if a low-energy plasma is used, or lower than 3  10 Pa if other triggering methods,
such as an energetic electron beam, UV ray, laser pulse, etc., are used. Care should be taken to choose a
power supply capable of reproducing the dynamic response of the array to transient short circuits (such as
limited overshoot current and fast recovery to the steady state). Simultaneous ESD current transient
monitoring and recording devices and a video imaging device are also required for the test.
7.3 External circuit
The cells need not be illuminated, but the available current and capacitance shall be simulated by power
sources and external capacitors, C and C , representing the satellite capacitance and solar array
sat CG
coverglass capacitance respectively. The capacitance of the missing coverglasses shall also be factored into
the test. The current waveform supplied by the external circuit shall be representative of the surface flashover
current in orbit. Under the present state of knowledge, the propagation distance is taken to be 2 m, confirmed
by a laboratory test using a 4 m  1 m coupon panel (see Reference [3]). The present best estimate of the
propagation speed of surface flashover is 10 km/s for a GEO solar array under inverted potential gradient.
See Annex F for an example of the current waveform derivation. The electric architecture of a solar paddle
shall be taken into account to determine the waveform. If a mitigation method for flashover current is included
in the design, the effect shall be taken into account in sizing energy in the capacitance C .
CG
The inductance of the wire harness should be representative of a flight solar panel. Excessive inductance
shall be avoided as it affects the transient current waveform. Short cable lengths or coaxial cables reduce the
unwanted inductance. Annex G specifies the external circuit for use in the case of inverted potential gradient.
If testing under normal potential gradient is imperative, a representative flashover current to the primary
discharge shall be provided.
12 © ISO 2011 – All rights reserved

7.4 CIC gap test — Test coupon and procedures
The test coupon(s) shall consist of at least two strings of two cells to represent a point surrounded by solar cell
corner edges. The total number of test coupon(s) shall reflect the production variation, including reworked
cells. The solar cells shall be laid down on the substrate in the same manner as for the flight model. The
substrate shall also be made of the same material as the flight model. If the solar panel design includes ESD
inception mitigation, whether the mitigation design is included in the test coupon or not depends on definition
of the worst case scenario in orbit. If failure of the mitigation method is regarded as the worst case scenario,
the mitigation method may be removed from the coupon. Before and after the test, the tasks specified in
Table 1 shall be carried out.
The primary discharges may be concentrated on the cell gap. The test shall ascertain that a significant
number of primary discharges (at least three, or more than three if agreed with the customer on the basis of
statistical discussion; see Reference [4] for an example) per given test condition occur in the active gap or in
the vicinity of the active gap if a grouted gap is tested.
Table 1 — Required tasks before and after the ESD test
Item Before test After test Comment
Visual inspection of the coupon x x With optical microscope
Output power measurement x x Same light source
Insulation check across the cell gap x x By measurement
Insulation check between the cells and
x x
the substrate
7.5 Panel test — Test coupon and procedures
The test coupon shall consist of at least three strings of three cells to represent means of cell surrounded by
other cells. The coupon shall be a flight-representative qualification coupon covering the production variation
of the string gap distance and CIC cell configuration (coverglass overhang, adhesive thickness, etc.), including
reworked cells. The solar cells shall be laid down on the substrate in the same manner as for the flight model.
The substrate shall also be made of the same material as the flight model. If the solar panel design includes
ESD inception mitigation, the mitigation design should be included in the test coupon so as to be as flight-
representative as possible. Before and after the test, the tasks specified in Table 1 shall be carried out.
The total number of primary discharges on the coupon shall be determined by means of statistical analysis.
The total number of ESD events, as derived in accordance with Annex D, is useful in determining the number
of primary discharges. No control shall be carried out at the primary discharge locations.
7.6 Success criteria
The test shall demonstrate that no damaging secondary arc occurs due to ESD.
8 Characterization tests for robustness to ESD and plasma interaction
8.1 Power degradation
8.1.1 Purpose
The purpose of this test is to characterize power degradation due to repeated ESD events after the desired
orbital lifetime. As the present knowledge about the power degradation is not mature enough to list power
degradation as a qualification item, the power degradation is measured for characterization purposes only.
This test may be performed simultaneously with the coupon panel test (7.5) provided the test requirements
listed in 7.2, 7.3 and 8.1 are satisfied.
8.1.2 Test facility
The test facility shall be the same as the one used for the ESD inception threshold test (6.2).
8.1.3 Test coupon
The CIC solar cells on the test coupon(s) shall be of flight quality. The definition of flight quality shall be
agreed upon with the customer prior to the test. The solar cell, coverglass, adhesive (both for coverglass and
substrate) and interconnector shall be of the same type and material as for the flight models. The illuminated
current-voltage (I-V) characteristics before the test shall be in accordance with the customer's specification of
flight quality. The substrate need not be of the same material as the flight model. The total number of test
coupon(s) shall reflect the production variation, including reworked cells.
8.1.4 External circuit
An external circuit shall provide additional electrostatic energy that would have been provided by the insulator
capacitance not accommodated inside the vacuum chamber. The current waveform supplied by the external
circuit shall be representative of the surface flashover current in orbit. Under the present state of knowledge,
the propagation distance is taken to be 2 m, confirmed by a laboratory test using a 4 m  1 m coupon panel
(see Reference [3]). The present best estimate of the propagation speed of surface flashover is 10 km/s for a
GEO solar array under inverted potential gradient. See Annex F for an example of the current waveform
derivation. The electric architecture of a solar paddle shall be taken into account in determining the waveform.
8.1.5 Test procedures
Illuminated current-voltage (I-V) characteristics shall be measured pre- and post-test with an appropriate light
source. The intensity of the light source shall be calibrated using a reference cell so that the relative change of
the power output can be measured precisely. The exact conditions of the illuminated I-V measurement shall
be based on an appropriate solar cell calibration standard, selected in agreement with the customer prior to
testing. A visual inspection shall be performed pre- and post-test using an optical microscope. The test can be
terminated once it satisfies one of the following:
a) the estimated number of ESD events during the mission in orbit (see 6.6) occurs;
b) ten discharges for any one cell do not create cell degradation;
c) a predetermined number of ESD events, as agreed with the programme customer, occurs.
It is advisable to measure dark current with a DC power supply in order to perform in situ monitoring of the
progress of degradation during the test by comparing the dark current with the pre-test value. See Annex H for
more information on the measurement of the dark current. The total drop of spacecraft power at the end of its
mission lifetime shall be derived and recorded.
14 © ISO 2011 – All rights reserved

8.2 Secondary arc
8.2.1 Purpose
The purpose of this test is to characterize the secondary arc for a given design of solar panel. The
characteristics include the probability of transition from a primary discharge to a secondary arc and the
duration of the secondary arcs if the transition probability is not zero. If the characteristics are already known
from past experience, this test is not necessary to qualify the solar panel for secondary arcs.
8.2.2 Triggering method and test facility
The triggering method and test facility shall be as specified in 7.2.
8.2.3 Test coupon
The test coupon(s) shall consist of at least two strings of two cells to represent a point surrounded by solar cell
corner edges. The total number of test coupon(s) shall reflect the production variation, including reworked
cells. The cell gap spacing should be kept within an allowable design tolerance from the specification. The
solar cells shall be laid down on the substrate in the same manner as for the flight model. The substrate shall
also be made of the same material as in the flight model.
If the solar panel design includes ESD inception mitigation, whether the mitigation design is included in the
test coupon or not depends on the definition of the worst case scenario in orbit. If failure of the mitigation
method is regarded as the worst case scenario, the mitigation method may be removed from the coupon.
8.2.4 External circuit
The external circuit shall be as specified in 7.3.
8.2.5 Test procedures
Before and after the test, the tasks specified in Table 1 shall be carried out. Primary discharges may be
concentrated on the cell gap. For each pair of gap voltage and string current agreed upon with the customer,
the probability of transition from primary discharge to secondary arc shall be derived. The test shall ascertain
that at least ten primary discharges per given test condition occur in the active gap (or in the vicinity of the
active gap in the case of a grouted gap). If it is necessary to identify the confidence level of the transition
probability, a significant number of primary discharges (as agreed with the customer on the basis of statistical
discussion) per the given test condition shall occur in the active gap. If the transition probability is not zero,
statistics of secondary arc duration, such as the mean and the standard deviation, are useful to extrapolate
the probability of a damaging secondary arc with a very long duration. If any statistical distribution function is
proposed to describe the secondary arc duration, its goodness-of-fit shall be determined by an appropriate
method, such as the  statistic. See Annex I for an example of deriving the statistics.
8.3 Power leakage to plasma
8.3.1 Purpose
The purpose of this test is to characterize the power leakage from a solar array to the surrounding plasma.
The plasma density and neutral gas density near the spacecraft shall be analysed first. Charging analysis
tools employing a particle simulation method (see Annex E) can estimate the amount of current for a given
plasma density. A rough estimate of the critical neutral gas density for transition of snapover to neutral gas
ionization can be obtained using formulas in References [5] and [6]. If the power leakage due to snapover is
low and neutral densities are too low for snapover to lead to neutral gas ionization, this test is not necessary.
8.3.2 Test facility
The test shall take place under vacuum in a test chamber equipped with a low-energy plasma source capable
of generating the plasma density expected around the solar array in orbit. The plasma species do not have to
be the same as those in orbit (see Reference [2]) A best effort shall be made to realize a plasma temperature
similar to the value in orbit. If the plasma temperature is higher than that in orbit, the test result on the
collected current shall be corrected based on either computer simulation or theory. The neutral background
pressure shall be kept below 0,01 Pa so that unrealistic electron-neutral collisions do not affect the test results.
The test facility shall be equipped with a DC power supply capable of supplying several amperes of current at
several hundred volts, a recording device to monitor the collection current and the coupon potential
simultaneously, a video imaging device and a plasma diagnostic device to determine the chamber plasma
properties including the plasma potential with respect to the chamber wall. See Annex C for the minimum
chamber size for a low-energy plasma test.
8.3.3 Test coupon
The number of CIC solar cells per coupon shall be as many as is practical. The minimum number shall be at
least three strings of three cells. The use of mechanical CIC cells (not electrically flight-like) is acceptable as
long as all the conductive parts, including the cell edge and interconnector, are exposed in the same manner
as the flight model. The substrate material shall be flight-representative. The total number of test coupon(s)
shall reflect the production variation including reworked cells.
8.3.4 External circuit
Both ends of the positive and negative electrodes of an array may be combined and connected to a DC power
supply as illustrated in Figure 5. The cells need not be illuminated during the test.

Key
1 coupon
2 plasma chamber
Figure 5 — Schematic of circuit diagram for the characterization test of plasma power leakage
16 © ISO 2011 – All rights reserved

8.3.5 Test procedures
Before the test, the solar array coupon shall be visually inspected with an optical microscope. The test coupon
bias voltage shall be changed step-wise while monitoring the potential of the coupon with respect to the
plasma potential until the potential reaches the maximum solar array output voltage in orbit. The current
collected at each potential shall be recorded after the current has become steady. After the test, the visual
inspection shall be repeated. The overall power leakage from the spacecraft to the plasma shall be derived
and reported.
8.4 Solar array back surface test
Although the scope of this International Standard is surface discharge and plasma interaction on the solar
array front surface, the possibility of ESD and plasma interaction on the solar array back surface cannot be
excluded (see Annex J).
9 Test report
The ESD test report shall include the following elements:
a) a title;
b) a description of the test facility (i.e. chamber size, chamber pressure, power supply, recording device,
etc.) with photographs of the test set-up; verify that all necessary facility criteria are met;
c) a description of the test coupon (i.e. solar cell type, solar cell output power, number of coupons tested,
etc.);
d) a description of the test conditions as follows:
1) when a low-temperature plasma is used, the plasma conditions, such as plasma density,
temperature and potential, shall be listed;
2) when an electron beam is used, the beam energy and current density shall be listed;
3) if the coupon is placed in a vacuum chamber for outgassing purposes prior to the test, the duration,
pressure and temperature shall be liste
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