SIST-TP CEN/TR 17603-20-06:2022
(Main)Space engineering - Assessment of space worst case charging handbook
Space engineering - Assessment of space worst case charging handbook
Common engineering practices involve the assessment, through computer simulation (with software like NASCAP [RD.4] or SPIS [RD.5]), of the levels of absolute and differential potentials reached by space systems in flight. This is usually made mandatory by customers and by standards for the orbits most at risk such as GEO or MEO and long transfers to GEO by, for example, electric propulsion.
The ECSS-E-ST-20-06 standard requires the assessment of spacecraft charging but it is not appropriate in a standard to explain how such an assessment is performed. It is the role of this document ECSS-E-HB-20-06, to explain in more detail important aspects of the charging process and to give guidance on how to carry out charging assessment by computer simulation.
The ECSS-E-ST-10-04 standard specifies many aspects of the space environment, including the plasma and radiation characteristics corresponding to worst cases for surface and internal charging. In this document the use of these environment descriptions in worst case simulations is described.
The emphasis in this document is on high level charging in natural environments. One aspect that is currently not addressed is the use of active sources e.g. for electric propulsion or spacecraft potential control. The tools to address this are still being developed and this area can be addressed in a later edition.
Raumfahrtproduktsicherung - Handbuch zu Minderungsmethoden von Strahlungseffekten auf ASICs und FPGA
Ingénierie spatiale - Guide sur les techniques de durcissement des ASICs et FPGAs vis-à-vis des effets des radiations
Vesoljski inženiring - Ocena priročnika za polnjenje v najslabšem primeru v vesolju
Običajne inženirske prakse vključujejo oceno ravni absolutnih in diferencialnih potencialov, ki jih dosegajo vesoljski sistemi med letom. Za oceno se uporabi računalniško simulacijo (s programsko opremo, kot sta NASCAP [RD.4] ali SPIS [RD.5]). To običajno zahtevajo stranke in standardi za orbite z največjim tveganjem, kot sta GEO ali MEO, in dolgi prenosi v GEO, na primer z električnim pogonom.
Standard ECSS-E-ST-20-06 zahteva oceno polnjenja vesoljskih plovil, vendar v njem ni pojasnjeno, kako se taka ocena izvaja. Naloga dokumenta ECSS-E-HB-20-06 je namreč, da podrobneje opiše pomembne vidike postopka polnjenja in poda napotke, kako izvesti oceno polnjenja z računalniško simulacijo.
Standard ECSS-E-ST-10-04 določa številne vidike vesoljskega okolja, vključno s plazemskimi in sevalnimi lastnostmi, ki ustrezajo najslabšim primerom površinskega in notranjega polnjenja. V tem dokumentu je opisana uporaba teh opisov okolja v simulacijah najslabšega primera.
Poudarek je na visoki ravni polnjenja v naravnem okolju. Eden od vidikov, ki trenutno ni obravnavan, je uporaba aktivnih virov, npr. za električni pogon ali nadzor potenciala vesoljskega plovila. Orodja za obravnavo tega se še razvijajo in to področje bo mogoče obravnavati v poznejši izdaji.
General Information
Standards Content (Sample)
SLOVENSKI STANDARD
01-marec-2022
Vesoljski inženiring - Ocena priročnika za polnjenje v najslabšem primeru v
vesolju
Space engineering - Assessment of space worst case charging handbook
Raumfahrtproduktsicherung - Handbuch zu Minderungsmethoden von
Strahlungseffekten auf ASICs und FPGA
Ingénierie spatiale - Guide sur les techniques de durcissement des ASICs et FPGAs vis-
à-vis des effets des radiations
Ta slovenski standard je istoveten z: CEN/TR 17603-20-06:2022
ICS:
49.140 Vesoljski sistemi in operacije Space systems and
operations
2003-01.Slovenski inštitut za standardizacijo. Razmnoževanje celote ali delov tega standarda ni dovoljeno.
TECHNICAL REPORT CEN/TR 17603-20-06
RAPPORT TECHNIQUE
TECHNISCHER BERICHT
January 2022
ICS 49.140
English version
Space engineering - Assessment of space worst case
charging handbook
Ingénierie spatiale - Guide sur les techniques de Raumfahrtproduktsicherung - Handbuch zu
durcissement des ASICs et FPGAs vis-à-vis des effets Minderungsmethoden von Strahlungseffekten auf
des radiations ASICs und FPGA
This Technical Report was approved by CEN on 29 November 2021. It has been drawn up by the Technical Committee
CEN/CLC/JTC 5.
CEN and CENELEC members are the national standards bodies and national electrotechnical committees of Austria, Belgium,
Bulgaria, Croatia, Cyprus, Czech Republic, Denmark, Estonia, Finland, France, Germany, Greece, Hungary, Iceland, Ireland, Italy,
Latvia, Lithuania, Luxembourg, Malta, Netherlands, Norway, Poland, Portugal, Republic of North Macedonia, Romania, Serbia,
Slovakia, Slovenia, Spain, Sweden, Switzerland, Turkey and United Kingdom.
CEN-CENELEC Management Centre:
Rue de la Science 23, B-1040 Brussels
© 2022 CEN/CENELEC All rights of exploitation in any form and by any means
Ref. No. CEN/TR 17603-20-06:2022 E
reserved worldwide for CEN national Members and for
CENELEC Members.
Table of contents
European Foreword . 6
Introduction . 7
1 Scope . 8
2 References . 9
3 Terms, definitions and abbreviated terms . 13
Terms from other documents . 13
Abbreviated terms. 13
4 Surface charging . 15
Fundamentals . 15
General methodology of surface charging analyses . 17
4.2.1 Introduction . 17
4.2.2 Necessity of 3D surface charging analyses . 17
4.2.3 Simulation process . 18
4.2.4 Assessment of simulation results . 19
Electrostatic discharge . 20
4.3.1 ESD types . 20
4.3.2 Thresholds for ESD occurrence . 20
4.3.3 Quantitative characterization of ESD electrical transients . 21
4.3.4 Interpretation of results . 25
Critical aspects with respect to worst case surface charging analyses . 25
4.4.1 Orbit . 25
4.4.2 Material properties . 26
4.4.3 Sunlit/Eclipse . 26
4.4.4 Protons . 27
4.4.5 Electric propulsion . 27
How to set up a simulation . 27
4.5.1 Charging environment parameters . 27
4.5.2 Modelling requirements for surface charging analyses . 27
4.5.3 Spacecraft geometry modelling . 28
4.5.4 Gmsh – The CAD interface to SPIS . 29
4.5.5 Physical groups and surface materials definition . 33
4.5.6 Basic electrical circuit of the satellite . 36
4.5.7 Plasma models . 37
4.5.8 Global parameters. 37
4.5.9 Consistency checks . 38
5 Internal Charging . 40
Fundamentals . 40
5.1.1 Introduction . 40
5.1.2 Floating metals . 40
5.1.3 Insulators . 40
5.1.4 Charge Deposition . 41
5.1.5 Conductivity . 41
5.1.6 Time-dependence . 43
General methodology . 43
5.2.1 Introduction . 43
5.2.2 Internal charging analyses . 44
5.2.3 Critical aspects with respect to worst case internal charging analysis . 45
5.2.4 Modelling aspects for internal charging analyses . 49
5.2.5 Environment . 50
5.2.6 Geometry . 50
5.2.7 Materials parameters . 51
5.2.8 Simulation tools in 1D and 3D . 51
5.2.9 Scenarios . 52
5.2.10 Important Outputs . 52
6 General aspects of surface and internal charging analysis . 53
Material characterization aspects . 53
Charging analyses and project phases . 53
6.2.1 Phase 0: Mission analysis . 53
6.2.2 Phase A: Feasibility . 53
6.2.3 Phase B: Preliminary definition . 53
6.2.4 Phase C: Detailed definition . 54
6.2.5 Phase D: Production . 54
6.2.6 Phase E: Utilisation . 54
Orbit plasma environment . 55
Plasma environment for different Earth orbits . 55
GEO worst case environments . 56
A.2.1 Introduction . 56
A.2.2 ECSS . 56
A.2.3 NASA . 56
A.2.4 ONERA/CNES . 58
LEO/Polar . 58
Figures
Figure 4-1: Current contributions influencing the surface charging of a body in space
plasma . 16
Figure 4-2: Flowchart showing the steps needed to determine the necessity of a 3D
surface charging analysis . 18
Figure 4-3: Flow diagram of the typical process of a 3D charging analysis . 19
Figure 4-4: Charged surface with area A showing the geometrical meaning and the
range for the parameter R . 23
Figure 4-5: Two-dimensional meshing of a solar array (from Sarrailh et al 2013 0) . 24
Figure 4-6: Examples of 2 discharges . 24
Figure 4-7: Definition of nodes and lines with Gmsh . 30
Figure 4-8: Definition of surfaces and volume with Gmsh . 31
Figure 4-9: Top: Surface meshes of the spacecraft and boundary. Bottom: Volume
mesh of the computational space. . 32
Figure 4-10: Definition of surface materials through the SPIS group editor . 33
Figure 4-11: Example of material properties list used by SPIS . 35
Figure 4-12: SPIS configuration of satellite electrical connections . 36
Figure 4-13: SPIS plasma parameters settings for the ECSS-E-ST-10-04 GEO worst
case environment for surface charging. . 37
Figure 5-1: Mulassis 0 simulation of net flux (forward minus backward travelling) due to
a 5 MeV incident beam in a planar sample of Aluminium. CSDA range is
approximately 11,4 mm . 41
Figure 5-2: Decision flow diagram for performing an internal charging analysis . 44
Figure 5-3: Current density v shielding depth curve for a geostationary orbit with
longitude 195deg East with nominal date 21/09/1994 according to the
FLUMIC model as calculated by the Mulassis tool in SPENVIS. The
FLUMIC spectrum was calculated by DICTAT in SPENVIS. . 46
Figure 5-4: Current density v shielding depth curve for the peak of the outer radiation
belt L=4,4, B/B0=1,0 with nominal date 21/09/1994 according to the
FLUMIC model as calculated by the Mulassis tool in SPENVIS. The
FLUMIC spectrum was calculated by DICTAT in SPENVIS. . 48
Tables
Table 4-1: Meanings of the material properties used in SPIS . 35
Table 5-1: Current density v shielding depth values for a geostationary orbit with
longitude 195deg East with nominal date 21/09/1994 according to the
FLUMIC model as calculated by the Mulassis tool in SPENVIS. The
FLUMIC spectrum was calculated by DICTAT in SPENVIS. . 47
Table 5-2: Current density v shielding depth values for the peak of the outer radiation
belt L=4,4, B/B0=1,0 with nominal date 21/09/1994 according to the
FLUMIC model as calculated by the Mulassis tool in SPENVIS. The
FLUMIC spectrum was calculated by DICTAT in SPENVIS. . 48
Table A-1 : Type of environments and order of magnitudes of density and temperature
encountered along typical orbits . 55
Table A-2 : Order of magnitudes of key plasma and charging parameters expected in
typical environments . 55
Table A-3 : ECSS-E-ST-10-04 worst case charging environment . 56
Table A-4 : NASA-HDBK-4002A worst case charging environment . 56
Table A-5 : NASA ‘more realistic’ geosynchronous worst case environment
specification . 57
Table A-6 : Severe charging environments in GEO 0 . 58
European Foreword
This document (CEN/TR 17603-20-06:2022) has been prepared by Technical Committee
CEN/CLC/JTC 5 “Space”, the secretariat of which is held by DIN.
It is highlighted that this technical report does not contain any requirement but only collection of data
or descriptions and guidelines about how to organize and perform the work in support of EN 16603-
20.
This Technical report (CEN/TR 17603-20-06:2021) originates from ECSS-E-HB-20-06A.
Attention is drawn to the possibility that some of the elements of this document may be the subject of
patent rights. CEN shall not be held responsible for identifying any or all such patent rights.
This document has been prepared under a mandate given to CEN by the European Commission and
the European Free Trade Association.
This document has been developed to cover specifically space systems and has therefore precedence
over any TR covering the same scope but with a wider domain of applicability (e.g.: aerospace).
Introduction
Spacecraft charging occurs due to the deposition of charge on spacecraft surfaces or in internal
materials due to charged particles from the environment. Resulting high voltages and high electric
fields cause electrostatic discharges which are a hazard to many spacecraft systems. Broadly speaking,
spacecraft charging can be divided into surface charging, which is caused by plasma particles with
energy up to several 10s of keV and internal charging which is caused by trapped radiation electrons
with energy around 0,2 MeV and above.
Both surface and internal charging have been associated with malfunctions and damage to spacecraft
systems over many years.
Scope
Common engineering practices involve the assessment, through computer simulation (with software
like NASCAP 0 or SPIS 0), of the levels of absolute and differential potentials reached by space
systems in flight. This is usually made mandatory by customers and by standards for the orbits most
at risk such as GEO or MEO and long transfers to GEO by, for example, electric propulsion.
The ECSS-E-ST-20-06 standard requires the assessment of spacecraft charging but it is not appropriate
in a standard to explain how such an assessment is performed. It is the role of this document ECSS-E-
HB-20-06, to explain in more detail important aspects of the charging process and to give guidance on
how to carry out charging assessment by computer simulation.
The ECSS-E-ST-10-04 standard specifies many aspects of the space environment, including the plasma
and radiation characteristics corresponding to worst cases for surface and internal charging. In this
document the use of these environment descriptions in worst case simulations is described.
The emphasis in this document is on high level charging in natural environments. One aspect that is
currently not addressed is the use of active sources e.g. for electric propulsion or spacecraft potential
control. The tools to address this are still being developed and this area can be addressed in a later
edition.
References
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environment
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charging
[RD.4] Myron J. Mandell, Victoria A. Davis, David L.
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[RD.5] Benoit Thiébault, Benjamin Jeanty-Ruard, Pierre
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Pierre Sarrailh, David Rodgers, Alain Hilgers,
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[RD.14] Insoo Jun, Henry B. Garrett, Wousik Kim, and
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[RD.16] Alex Hands, Keith Ryden, Craig Underwood, David
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(MOBE-DIC) IEEE TRANSACTIONS ON NUCLEAR
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[RD.17] G. P. Ginet, P. O’Brien, S. L. HustonW. R. Johnston,
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[RD.19] D. Payan, A. Sicard-Piet, J.C. Mateo-Velez, D.Lazaro,
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[RD.22] NASA-HDBK-4002A, Mitigating in space charging
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[RD.32] Sarno-Smith, Lois K., Larsen, Brian A., Skoug, Ruth
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[RD.34] NASA Technical paper 2361,1984 Design guidelines
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[RD.35] Matéo-Vélez, J.-C., Sicard, A., Payan, D.,
Ganushkina, N., Meredith, N. P., & Sillanpäa, I.
(2018). Spacecraft surface charging induced by
severe environments at geosynchronous orbit. Space
Weather, 16. https://doi.org/10.1002/2017SW001689
Terms, definitions and abbreviated terms
Terms from other documents
a. For the purpose of this document, the terms and definitions from ECSS-S-ST-00-01 apply, in
particular the following terms:
1. environment
Abbreviated terms
For the purpose of this document, the abbreviated terms from ECSS-S-ST-00-01 apply and in
particular the following:
Abbreviation Meaning
astronomical unit
AU
beginning-of-life
BOL
computer-aided design
CAD
continuous slowing down approximation
CSDA
(relating to range of radiation in matter)
electromagnetic compatibility
EMC
electrostatic discharge
ESD
end-of-life
EOL
Geometry Definition Markup Language
GDML
graphical user interface
GUI
geostationary orbit
GEO
geostationary transfer orbit
GTO
medium Earth orbit
MEO
multi-layer insulation
MLI
LoaSAlamos National Laboratory
LANL
low Earth orbit
LEO
secondary electron emission
SEE
In addition the following acronyms are used in this document
Acronym Meaning
United States Airforce electron environment model / Standard
AE9/SPM
Plasma Model
Software with ability to perform 3D internal charging
CIRSOS
simulations (Collaborative Iterative Radiation Simulation
Optimisation Software)
Software for 1D Internal charging analysis (Dielectric Internal
DICTAT
Charging Threat Assessment Tool)
An ESA study of Solar Array Triggering Arc Phenomena
EMAGS-3
A worst case electron environment model for internal
FLUMIC
charging (Fluence Model for Internal Charging)
Software generator of finite element meshes
Gmsh
Software for 1D Internal charging analysis (Monte Carlo
MCICT
Internal Charging Tool)
A worst case electron environment model for internal
MOBE-DIC
charging (Model of Outer Belt Electrons for Dielectric
Internal Charging)
Software for 3D simulation of surface charging (NASA
NASCAP
Charging Analysis Program)
Software for 1D Internal charging analysis (Numerical
NUMIT
Integration)
A spacecraft dedicated to surface charging observations
SCATHA
(Spacecraft Charging at High Altitudes)
Software for 3D simulation of surface charging (Spacecraft
SPIS
Plasma Interaction Simulation)
Web service for space environment analysis (Space
SPENVIS
Environment Information System)
Surface charging
Fundamentals
This Section gives a brief overview of the most important physical mechanisms connected with the
surface charging of a body which is exposed to the space environment. Some of the randomly
propagating charged particles incident on the surface of the body are collected by the surface, while
the others are deflected or re-emitted. The ratio between these two contributions depends on the
surface material and on the energy and angle of incidence of the impacting particles. The collected
particles can be seen as a current from the plasma environment to the satellite. Because negatively
charged electrons have a higher mean velocity than positively charged ions the spacecraft usually
tends to assume absolute negative potentials.
Besides the simple collection of charged particles, additional effects can lead to a
re-emitted current from the surface. So, incoming particles are able to eject electrons from the surface
material. This effect is called secondary electron emission (SEE). This process is energy dependent and
for some materials the ratio of released electrons to incident electrons (SEE yield) can be greater than
one. Depending on the considered material an SEE yield of more than five is possible. Generally, there
is a peak in secondary electron emission between several 100 eV and a few keV. Ambient electrons
with energy between the two crossover points (the two energies where the yield is 1) in the yield
curve, tend to charge the spacecraft positive instead of negative. In addition to SEE, the UV
component of sunlight is also able to release electrons from the surface in what is called the
photoelectric effect. Both SEE and photoemission play an important role in limiting the negative
charging of the corresponding spacecraft surfaces.
Besides these direct currents to and from the surface to the plasma there are currents between
different surfaces through the grounding network of the satellite. Note that a small current can even
bleed through the thin dielectric (insulating) surfaces and this can have an impact on the final current
balance on these surfaces.
The charging process continues until an equilibrium condition is reached, which is characterized by a
balance between all collected and emitted currents (positive and negative). In Figure 4-1 the main
current contributions connected with the charging of a body in space plasma are depicted. At
equilibrium the black and the blue currents are balanced.
The charging process is influenced by a lot of different physical effects defining and changing the
surface currents as explained above. Additionally, the currents arising from the external plasma are
changed by the spacecraft potentials which extend out into the plasma. Also currents flow from one
part of the spacecraft to another through the grounding network.
Hence, when a realistic assessment of the charging threat on a satellite is performed, the use of a
dedicated 3D simulation tool where all these effects are taken into account is essential. Note that there
can be large differences between 1D/2D tools and a fully 3D simulation due to the influence of the
charged spacecraft on the plasma creating sheaths which can only be considered in a 3D tool.
solar radiation
photoemission
incident ions
SEE from electrons and ions
incident electrons
backscattered electrons
leakage current
Figure 4-1: Current contributions influencing the surface charging of a body in
space plasma
Surface charging and the corresponding effects and possible threats to the spacecraft functionality can
be divided into two major processes:
• Absolute charging – this relates to the changing floating potential of the grounded parts of the
satellite with respect to the ambient plasma environment. Usually, the time to reach equilibrium
floating potential is very short, i.e. some milliseconds, because of the very low capacitance of
the spacecraft w.r.t. the plasma environment. However, depending on the material distribution
on the satellite this process can sometimes last longer, i.e. minutes, especially if the material
distribution is dominated by dielectric materials.
• Differential charging - because of different material properties in terms of SEE and
photoemission yields, different materials on the surface of the spacecraft tend to charge
unequally. Differential charging can even occur between surfaces made of identical materials
where one surface is sunlit and the other is in shade, or because of different potential barriers
above these surfaces forcing secondary electrons to be recollected locally. For differential
charging, it can take several minutes or hours to reach equilibrium. The timescale is ruled by
the relatively high capacitances of spacecraft surface areas w.r.t. the grounded layers beneath.
Differential charging is reduced by current leakage across the surface and through the dielectric
thickness.
In eclipse conditions and for a low energy plasma equilibrium potentials in volts are usually on the
order of two to three times the plasma energy in eV. These plasma conditions are usually encountered
in the ionosphere and plasmasphere at low altitudes with low inclination below 50 degrees and are
usually considered uncritical. In a very dense and low energy plasma the satellite absolute potential
can be driven by the voltage of the space exposed interconnectors on the solar array. These are often
positively biased with respect to spacecraft ground and attract electrons so that the satellite absolute
potential becomes negative.
For LEO orbits with a high inclination, such as sun-synchronous orbits, the satellite has crosses the
auroral zone where higher energy (~10 keV) electrons can be found. In this region, are found auroral
arcs where high energy electrons are streaming towards lower altitudes due to geomagnetic storms.
Simultaneous with the arrival of these high energy electrons the cold background plasma density can
also be greatly reduced in this region, producing a hazardous environment with a chance of strong
satellite charging. In these auroral charging conditions the satellite potential can drop down to levels
around -1kV with differential potentials that can also be rather high so that electrostatic discharges
(ESDs) are possible.
Orbits around the earth above the ionosphere and plasmasphere are mainly characterised by plasma
with a small total density. The energy distribution of this plasma varies as a result of changing
geomagnetic configurations and due to the injection and subsequent decay of hot plasma during
geomagnetic substorms. Hence the range of possible potentials in the outer magnetosphere is also
wide. When secondary and photoemission effects dominate, the satellite potential is close to 0 V but
for worst case conditions with a large amount of electrons at high energies, well above the second
cross-over point in the yield curves, charging to very high negative potentials, in the range of -10 kV,
is possible. If insulating satellites surfaces are used there can also be a high potential difference
between sunlit and shaded surfaces due to the photoemission which cannot be balanced between
insulating surfaces.
General methodology of surface charging analyses
4.2.1 Introduction
Since there are many aspects of the space environment that can have a detrimental effect on a
spacecraft or its operations, it is part of the spacecraft design process to make an assessment of how
severe the effects can possibly be. These analyses require an understanding of the environment and
can involve testing in the laboratory, computer simulation or both. Where space systems are re-used
in the same or a similar environment, heritage can also be relevant
Spacecraft surface charging has a strong dependence on the complete spacecraft geometry. Although
individual exposed structures and materials are often tested in the laboratory under charged particle
irradiation, realistic laboratory testing of a whole spacecraft is seldom possible and so computer
simulation plays a very important role.
In this Section we discuss the general requirements and methods of surface charging analyses.
4.2.2 Necessity of 3D surface charging analyses
The first decision which needs to be taken in a new project is on the necessity of a dedicated 3D
surface charging analysis. The flowchart given in Figure 4-2 shows the corresponding decision tree
which is usually followed to determine if a 3D analysis is performed. So, it is important to consider
whether the satellite is launched into a critical orbit (GEO, MEO, GTO, Polar) where surface charging
and possible ESD can occur. In such a case it is strongly recommended to perform a charging analysis
to show the robustness of the satellite design with respect to the plasma environment. The same is
recommended if special requirements with respect to charging are applicable for the mission, e.g.
scientific missions with requirements for the maximum allowable differential potential on the satellite.
Satellite launched into Special Requirements
No
critical orbit for surface potentials
No
Yes
Perform 3D Charging
Yes Assessment using 1D tools
Analysis
Figure 4-2: Flowchart showing the steps needed to determine the necessity of a 3D
surface charging analysis
4.2.3 Simulation process
The typical analysis process including the necessary inputs and outputs and interpretation of results is
shown in Figure 4-3. The main inputs for the 3D analysis are the basic electrical circuit of the satellite
along with a dedicated 3D CAD model with the most important space exposed surface materials
included. The plasma environment definition coming from the project specific environmental
specification or the applicable standard (usually ECSS-E-ST-10-04 0) is also an important input to the
analysis.
3D Geometry Model
Basic Electrical Circuit of the Plasma Environment
including representative
Satellite Definition
surface material distribution
3D Charging Simulation
Assessment of surface
3D Surface Potential material change for
Distribution on the satellite improved charging
behaviour
Potential Gradient
No Yes
Thresholds Violated
No
Acceptable wrt.
Satellite design safe wrt
Yes Requirements and for
surface charging
customer
Figure 4-3: Flow diagram of the typical process of a 3D charging analysis
4.2.4 Assessment of simulation results
The main output of the analysis is the surface potential distribution on the 3D model of the satellite
which can then be used to determine the potential differences to be expected on the satellite.
Maximum differential potentials for dielectrics with respect to ground or an adjacent metal are then
compared against the generic requirements for the onset of ESD given in ECSS-E-ST-20-06:
• Where the conductor is positive with respect to the dielectric (‘normal potential gradient’)
• Where the conductor is negative with respect to the dielectric (‘inverted potential gradient’)
When special requirements which are more stringent than the generic ones, for the satellite are
applicable then it is important to use these requirements in the assessment of the results. If the
applicable thresholds are not violated the analysis is finished and the satellite design with respect to
surface charging can be consolidated. If the thresholds are violated there are in principle two options:
a. The design of the satellite can be changed in order to get a better charging behaviour e.g. by
using more conductive materials or changing to materials with higher SEE yield. The effectivity
of the proposed design changes is then verified by an update of the charging analysis taking the
changes into account.
b. If a design change is not possible or easily done there is also the possibility to prove that the
violation of the requirements poses no major threat or impediment to the satellite mission. In
this case the ECSS procedure requires that the customer is involved in the acceptance of such a
non–compliance.
Electrostatic discharge
4.3.1 ESD types
There are different basic types of ESD, depending on the materials which are involved in the event.
• ‘Metal ESD’ is a direct discharge between two conducting parts of the structure. Due to the very
good conductivity of the materials involved the ESD pulse is very fast discharging the elements,
which leads to very high peak currents. The very high current can cause permanent damage to
the involved materials and in combination with the short pulse duration strong electromagnetic
interference can be created which can disturb other systems on the satellite. Due to these
dangerous properties metal ESDs are generally prevented on a real system by means of
grounding of all conducting parts to the structure.
• A ‘triple point’ where metal, dielectric and vacuum come together is particularly associated
with strong ESDs. It can become a discharge site when an inverted potential gradient (IPG)
arises, i.e. where a conductor is more negative than an insulator. A strong electric field can exist
between the negative metal and the less-negative dielectric. Electric field-induced electron
emission is initiated from an irregularity on the metal surface and the electrons impact the
adjacent dielectric. If the impact energy is sufficient, there is secondary emission which drives
the dielectric positive and increases the electric field between metal and dielectric which further
increases the electron emission from the metal in a run-away process. Triple point discharges
can lead to ‘flash-over’ (described in Section 4.3.3)
• Other discharges are dielectric material discharges (as described in ECSS-E-ST-20-06) which
include surface propagating discharges that can be triggered on dielectric coatings by punch-
through of the dielectric.
4.3.2 Thresholds for ESD occurrence
Depending on the exact situation there are mainly three different threshold values defined for the
onset of ESDs:
• The straightforward value is the material specific dielectric breakdown field strength. If this
threshold is reached the charge stored on top of an insulator is released to the conducting layer
beneath. This is a discharge event where the transient goes through the material. If the exact
value for the dielectric breakdown voltage or field strength is not exactly known, ECSS-E-ST-20-
06 clause 6.2.1 gives a default value to be used.
• Two other types of discharges involve dielectric surfaces:
Direct potential gradient ESD, where the dielectric surface is at a more negative potential
than an adjacent conducting one. In this case of the threshold for the risk of ESD is given
by ECSS-E-ST-20-06 clause 6.2.1.
Inverted potential gradient (IPG) where the conductor is at a more negative potential
compared to the insulator. In this case the threshold value is lower because field emission
of electrons from the conductor plays an important role in the triggering of the ESD.
Concerning the exact value for the onset of ESDs in this scenario the literature values
differ. The ECSS-E-ST-20-06 standard gives one value [clause 6.2.1], whereas NASA
literature 0 sta
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