ISO 19923:2017

Space environment (natural and artificial) — Plasma environments for generation of worst case electrical potential differences for spacecraft

ISO 19923:2017 specifies space plasma environments that lead to the generation of the worst-case surface potential differences for spacecraft. It also specifies how to estimate worst-case potential differences by using the simulation codes provided. ISO 19923:2017 includes plasma energy and density in GEO, PEO, and MEO. ISO 19923:2017 does not include descriptions of plasma energy and density in LEO because large surface charging in LEO is likely to be due to high-voltage power generation by instrumentation of the spacecraft. ISO 19923:2017 deals with external surface charging of spacecraft only.

Environnement spatial (naturel et artificiel) — Environnements plasmatiques pour la génération de différences de potentiel électrique les plus défavorables pour les véhicules spatiaux

General Information

• Status: Published
• Publication Date: 13-Jun-2017
• ICS: 49.140 - Space systems and operations
• Technical Committee: ISO/TC 20/SC 14 - Space systems and operations
• Drafting Committee: ISO/TC 20/SC 14 - Space systems and operations
• Current Stage: 9060 - Close of review
• Start Date: 02-Dec-2027

Standards Content (Sample)

INTERNATIONAL ISO
STANDARD 19923
First edition
2017-06
Space environment (natural and
artificial) — Plasma environments
for generation of worst case electrical
potential differences for spacecraft
Environnement spatial (naturel et artificiel) — Environnements
plasmatiques pour la génération de différences de potentiel électrique
les plus défavorables pour les véhicules spatiaux
Reference number
ISO 19923:2017(E)
©
ISO 2017

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ISO 19923:2017(E)

COPYRIGHT PROTECTED DOCUMENT
© ISO 2017, Published in Switzerland
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ii © ISO 2017 – All rights reserved

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ISO 19923:2017(E)

Contents  Page
Foreword .iv
1 Scope . 1
2  Normative references . 1
3  Terms and definitions . 1
4  Symbols and abbreviated terms . 2
5  Criteria for worst-case environment . 2
6  Procedures for application to spacecraft design . 2
7  Space environments for worst-case simulations . 3
7.1 GEO worst-case environment. 3
7.2 PEO and MEO worst-case environments . 3
Annex A (informative) Spacecraft charging analysis tools. 4
[7]
Annex B (informative) Round-robin simulation . 5
Annex C (normative) Material properties .10
Annex D (informative) Tailoring guideline for this document .13
Bibliography .14
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ISO 19923:2017(E)

Foreword
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This document was prepared by Technical Committee ISO/TC 20, Aircraft and space vehicles,
Subcommittee SC 14, Space systems and operations.
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INTERNATIONAL STANDARD  ISO 19923:2017(E)
Space environment (natural and artificial) — Plasma
environments for generation of worst case electrical
potential differences for spacecraft
1 Scope
This document specifies space plasma environments that lead to the generation of the worst-case
surface potential differences for spacecraft. It also specifies how to estimate worst-case potential
differences by using the simulation codes provided.
This document includes plasma energy and density in GEO, PEO, and MEO. This document does not
include descriptions of plasma energy and density in LEO because large surface charging in LEO is
likely to be due to high-voltage power generation by instrumentation of the spacecraft.
This document deals with external surface charging of spacecraft only.
2  Normative references
There are no normative references in this document.
3  Terms and definitions
For the purposes of this document, the following terms and definitions apply.
ISO and IEC maintain terminological databases for use in standardization at the following addresses:
— IEC Electropedia: available at http:// www .electropedia .org/
— ISO Online browsing platform: available at http:// www .iso .org/ obp
3.1
double Maxwellian distribution
electron and proton distribution functions in GEO fitted with two temperatures
[12]
Note 1 to entry: Maxwellian distribution is as follows :
32/
 2 2 
   
n n
m mv mv
 
1 2
 
fv()=−expe + xxp− 
 
32//  32  
22p kT 2kT
 
 
1 2
()kT ()kT
   
 1 2 
where
m is the mass of particle;
−23
k is the Boltzmann constant 1,380 648 52 × 10 J/K;
n , n are the number density of particle;
1 2
T , T are the temperature of particle.
1 2
3.2
differential voltage
differential potential
potential difference between any two points in spacecraft, especially the insulator surface and the
spacecraft body, during differential charging
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ISO 19923:2017(E)

3.3
inverted potential gradient
result of differential charging where the insulating surface or dielectric reaches a positive potential with
respect to the neighbouring conducting surface or metal: PDNM (positive dielectric negative metal)
3.4
normal potential gradient
result of differential charging where the insulating surface or dielectric reaches a negative potential
with respect to the neighbouring conducting surface or metal: NDPM (negative dielectric positive metal)
3.5
surface charging
deposition onto or the removal of electrical charges from external surfaces of the spacecraft
4  Symbols and abbreviated terms
−19
eV electron volt, where 1 eV = 1,602 × 10 J
GEO geosynchronous orbit
LEO low Earth orbit
MEO medium Earth orbit
PEO polar Earth orbit
Ne electron density
Ni ion density
Te electron temperature
Ti ion temperature
5  Criteria for worst-case environment
The worst-case environment shall be defined as the space environment measured in space that causes
the maximum potential difference between the spacecraft electrical grounding body and external non-
conductive surfaces or isolated conductive surfaces. Worst-case conditions shall be realistic.
Combinations of densities and temperatures for a valid worst-case condition shall be subject to all of
the following:
— reported in the literature or published databases;
— checked to make sure they are based on valid measurements;
— physically realistic (i.e. do not violate energy density or other physical requirements); and
— verified using good spacecraft charging codes (i.e. COULOMB-2, MUSCAT, SPIS, NASCAP-2k).
This document is a part of spacecraft charging design.
6  Procedures for application to spacecraft design
Spacecraft charging simulation should be carried out at an early stage of spacecraft design. Ideally, this
should be before selecting the materials for those spacecraft surfaces that will be exposed to the space
environment.
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ISO 19923:2017(E)

Use worst-case environments mentioned in Clause 7 as input parameters for charging simulations.
Material properties for spacecraft charging can change after exposure to the space environment. If
possible, employ simulation tools using material properties after the appropriate space environmental
ageing. See Annex C.
Radiation induced conductivity can change the bulk resistivity of materials. If possible, employ
simulation tools that use the material properties after exposure and ageing in the appropriate space
[11]
environment .
In the computer simulations, use the appropriate spacecraft geometry, material data, and environmental
conditions. Run the simulation from a zero charging initial condition until differential potentials fully
develop.
For examples of simulation codes, see Annex A. Note, however, that the list of codes in Annex A is not
exclusive.
7  Space environments for worst-case simulations
7.1  GEO worst-case environment
The double Maxwellian distribution contained in Table 1 shall be used for worst-case simulation.
Table 1 — Space environment cases simulated
Ne1 Te1 Ne2 Te2 Ni1 Ti1 Ni2 Ti2
−3 −3 −3 −3
m eV m eV m eV m eV
2,00E+05 400 2,30E+06 24 800 1,60E+06 300 1,30E+06 28 200
Other worst cases have been proposed. See Annex B for comparisons. m and m are
e i
−31 −27
9,109 383 56 × 10 kg and 1,672 621 9 × 10 kg, respectively.
7.2  PEO and MEO worst-case environments
The worst-case plasma environment in PEO and MEO will be updated as more published measured
environments become available. See Reference [3] for one published PEO environment.
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ISO 19923:2017(E)

Annex A
(informative)

Spacecraft charging analysis tools
A.1 COULOMB-2
[4]
COULOMB-2 code can be applied to modelling of spacecraft charging in PEO and GEO. For building
of the spacecraft geometrical models and modelling results visualization, the SALOME platform is
used. Plasma currents are computed in terms of Langmuir equations and particle trajectory modelling.
Integral equation method is used for electrostatic equation solving. Database of electro-physical
properties of typical spacecraft materials is also included in the code. The code is not easily available
outside Russia.
A.2 MUSCAT
[5]
MUSCAT is a fully 3D particle code that can be applied to spacecraft in LEO, PEO and GEO. Its
algorithm is a combination of PIC and particle tracking. A parallel computation technique is used
for fast computation. It has a JAVA-3D based graphical user interface for 3D modelling of spacecraft
geometry and output visualization. The surface interactions included in the NASCAP series and SPIS
are modelled. A material property database is also included. The code is commercially available.
A.3 NASCAP-2k
The most recent NASCAP code (NASCAP-2k) is available, free, to US citizens only. This is a comprehensive
code with realistic geometry. It is reported to combine the capabilities of NASCAP-GEO, NASCAP-LEO
and POLAR. The code is not easily available outside the US.
A.4 SPIS
[6]
SPIS is a fully 3D PIC code that allows the exact computation of the sheath structure and the current
collected by spacecraft surfaces for very detailed geometries. Surface interactions including photo-
electron emission, back-scattering, secondary-electron emission and conduction are modelled. The
source code is freely available from www .spis .org and a mailing list provides a limited amount of
support.
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ISO 19923:2017(E)

Annex B
(informative)

[7]
Round-robin simulation
B.1 Round-robin simulations with NASCAP-2k
In order to estimate the degree of charging on spacecraft in GEO charging environments, a generic
spacecraft model was constructed. It is shown in Figure B.1. The back sides of the arrays were covered
with graphite. Dimensions of the model are the following.
The body is X: 1,86 m; Y: 1,55 m; Z: 2,56 m. The NPaint box on the top is X: 0,62 m; Y: 0,516 m; Z: 0,62 m.
The aluminium box at the bottom is X: 0,30 m; Y: 1,55 m; Z: 0,62 m. The solar arrays have a width of
2,5 m; length: 4,0 m; thickness: 0,10 m; twist: 45 degrees. The solar array booms are 2,0 m long and
0,10 m square in cross-section. The round antenna is 2,5 m in diameter and separated from the body by
0,3 m. Material properties are shown in Table B.1.
Figure B.1 — Calculation model with NASCAP-2k
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ISO 19923:2017(E)

Table B.1 — Material properties
Coverglass  Dielec- Thick- Bulk Atomic δ E Proton  Proton Photoemis- Surface Atomic Densi-
max max
material tric ness conductiv- number keV yield max sion resistiv- wt ty
−2 −3
constant m ity eV A m ity amu kg m
−1 −1
Ω m Ω/square
Graphite 1 1,00E-03 −1 4,5 0,93 0,28 0,455 80 7,20E-06 −1 12,01 2 250
Aluminium 1 1,00E-03 −1 13 0,97 0,3 0,244 230 4,00E-05 −1 26,98 2 699
Black
3,5 2,50E-06 −1 5 5,2 0,90 0,455 140 5,00E-06 −1 12,01 1 600
a
Kapton®
a
Kapton® 3,5 1,27E-04 1,00E-16 5 2,1 0,15 0,455 140 2,00E-05 1,00E+16 12,01 1 600
Solar cells
3,8 1,25E-04 1,00E-13 10 5,8 1 0,244 230 2,00E-05 1,00E+19 20 2 660
(MgF2)
OSR 4,8 1,50E-04 1,00E-16 10 3,3 0,5 0,455 140 2,00E-05 1,00E+19 20 2 660
NPaint 3,5 1,27E-04 1,00E-16 5 2,1 0,15 0,455 140 2,00E-05 1,00E+16 12,01 1 600
a
  Kapton® is the trade name of a product supplied by DuPont. This information is
...