on-orbit detection of spacecraft charging effects
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On-Orbit Detection of Spacecraft Charging Effects
Joseph I. MinowNASA, Marshall Space Flight Center
In-Space Inspection Workshop 2017NASA JSC, Houston, Texas
30 January – 2 February 2017
ISS image: 7 March 2012 1
Introduction
Anomaly Diagnosis Number %
ESD-Internal, surface 162 54.1
and uncategorized
SEU (GCR, SPE, radiation belt) 85 28.4
Radiation dose 16 5.4
Meteoroids, orbital 10 3.3
debris
Atomic oxygen 1 0.3
Atmospheric drag 1 0.3
Other 24 8.0
Total 299 100.0%
Space Environment Impacts on Space Systems [Koons et al., 2000]*
*Sources: Aerospace Corporation study using data from1) NOAA/NGDC Spacecraft Anomaly Manager2) NASA Anomaly Reports3) USAF 55th Space Weather Squadron anomaly database4) Individual Program Offices databases
2
Introduction
Outline
• Introduction to spacecraft charging
– Why is it important
– What causes it?
– Where it is important?
• Impacts to spacecraft
• Examples of charging damage
• Sensors
• Opportunities for in-space inspection
Anomaly Diagnosis Number %
ESD-Internal, surface 162 54.1
and uncategorized
SEU (GCR, SPE, radiation belt) 85 28.4
Radiation dose 16 5.4
Meteoroids, orbital 10 3.3
debris
Atomic oxygen 1 0.3
Atmospheric drag 1 0.3
Other 24 8.0
Total 299 100.0%
Space Environment Impacts on Space Systems [Koons et al., 2000]*
*Sources: Aerospace Corporation study using data from1) NOAA/NGDC Spacecraft Anomaly Manager2) NASA Anomaly Reports3) USAF 55th Space Weather Squadron anomaly database4) Individual Program Offices databases
3
Anomalies and Failures Attributed to Charging
Spacecraft Year(s) Orbit Impact* Spacecraft Year(s) Orbit Impact*
DSCS II 1973 GEO LOM Intelsat K 1994 GEO Anom
Voyager 1 1979 Jupiter Anom DMSP F13 1995 LEO Anom
SCATHA 1982 GEO Anom Telstar 401 1994, 1997
GEO Anom/LOM
GOES 4 1982 GEO LOM TSS-1R 1996 LEO Failure
AUSSAT-A1, -A2, -A3 1986-1990 GEO Anom TDRS F-1 1986-1988 GEO Anom
FLTSATCOM 6071 1987 GEO Anom TDRS F-3,F-4 1998-1989 GEO Anom
GOES 7 1987-1989 GEO Anom/SF INSAT 2 1997 GEO Anom/LOM
Feng Yun 1A 1988 LEO Anom/LOM Tempo-2 1997 GEO LOM
MOP-1, -2 1989-1994 GEO Anom PAS-6 1997 GEO LOM
GMS-4 1991 GEO Anom Feng Yun 1C 1999 LEO Anom
BS-3A 1990 GEO Anom Landsat 7 1999-2003 LEO Anom
MARECS A 1991 GEO LOM ADEOS-II 2003 LEO LOM
Anik E1 1991 GEO Anom/LOM TC-1,2 2004 ~2xGTO, GTO Anom
Anik E2 1991 GEO Anom Galaxy 15 2010 GEO Anom
Intelsat 511 1995 GEO Anom Echostar 129 2011 GEO Anom
SAMPEX 1992-2001 LEO Anom Suomi NPP 2011-2014 LEO Anom
*Anom=anomaly, SF=system failure, LOM=Loss of mission 4
Surface charging
Internal (deep dielectric) charging
Electrodynamic potentials
Electric Potentials on Spacecraft Surfaces
Electrostatic potentials• Electrostatic potentials generated by net
charge density on spacecraft surfaces or within materials due to current collection to/from the space environment
• Examples include– Plasma currents to surfaces– Secondary electron currents– Photoelectron currents– Solar array current collection– Conduction currents– Active current sources (Electron, ion
beams, electric thrusters, plasma contactors)
Electrodynamic (inductive) potentials• Modification of spacecraft frame potential
distribution without change in net charge• External plasma environment not required• Examples include
– EMF generated by motion of conductor through magnetic field
– Externally applied electric fields
2
)(
ED
C
CC
m
SdBvE
SdBvESdE
BvEE
EqF
BvEqF
)(
)(
'
)(
Laboratory frame
Spacecraft rest frame
Forces equal in both frames!
[c.f., Whipple, 1981; p. 272 Wangness, 1986; p. 210 Jackson, 1975; Maynard, 1998]
k
kIdt
dC
dt
dQ ~ 0 at equilibrium
5
Surface Charging Physics
• Surface charging is a current balance process to and from spacecraft surfaces as a function of the spacecraft potential
k
kIAdt
dσ
dt
dVC
dt
dQ
(V)I
(V)I
(V)I
(V)I
(V)I
(V)I
(V)I
Idt
dQ
eph,
si
se
c
ebs,
e
i
k
k
incident ions
incident electrons
backscattered electrons
conduction currents
secondary electrons due to Ie
secondary electrons due to Ii
photoelectrons
(Garrett and Minow, 2004)
few eV to ~50 keV
6
7
Internal Charging Physics
σERJCJRJJ
Jt
ρ
0κεε εE,D
ρD
~50 keV to few MeV
1.0α0.5α
dt
dγk
radiationσ
Eradiation
σdark
σRJ
Threat Environment: Outer Radiation Belt
Outer radiation belt
• Source: hot magnetospheric plasma during geomagnetic storms, trapped outer radiation belt electrons
• Threat orbits:
– GEO (6.6 Re)
– GTO and HEO (LEO to GEO)
– GPS (4.1 Re)
– Earth escape radiation belt transit
• Types:
– Surface charging (10’s keV storm plasma)
– Internal charging (MeV outer radiation belt electrons)
8
Image: JHU/APL, NASA
Image: NOAA
Threat Environment: LEO Polar, SAA
Low Earth orbit, polar regions
• Source: auroral electrons, trapped inner radiation belt electrons in South Atlantic Anomaly
• Threat orbits:
– 100’s km to 10,000’s km altitude
– Inclinations greater than about 50
• Types:
– Surface charging (10’s keV auroral electrons)
– Internal charging (100’s keV inner radiation belt electrons)
9
NOAA SWPC
NOAA SWPC NOAA SWPC
>300 keV>30 keV
GEO Surface Charging and Electron Temperature
• Significant negative charging doesn’t develop until enough electrons in environment have energies greater than the second cross over energy
• Hot electrons in the space plasma environment are required for strong negative charging
ATS-6X SCATHA
LANL
(Lai et al, 2003)10
[Olsen, 1983]
• Charging anomalies and failures depend on– Magnitude and gradients in the induced potentials and strength of the electric fields
– Material configuration (and capacitance)
– Electrical properties of the materials
• Surface and volume resistivity, dielectric constant
• Secondary and backscattered electron yields, photoemission yields
• Dielectric breakdown strength
Charging Anomaly and Failure Mechanism
• Accumulation of excess negative charge (or inductive charge redistribution) generates potential differences between spacecraft and space (frame potential) or between two points on the spacecraft (differential potential)
• Potential gradient produces an electric field E = -
• An electrostatic discharge (ESD) can occur when the electric fields associated with potential differences exceed the dielectric breakdown strength of materials allowing charge to flow in an arc current
• Arc currents deposit energy due to Joule heating, damage depends on energy available to arc
Energy = ½CV2 PMMA (acrylic) charged by ~2 to 5 MeV electrons
inches
11
Impact of Charging on Spacecraft Systems
Electrostatic discharge (ESD) currents
• Compromised function and/or catastrophic destruction of sensitive electronics
• Solar array string damage (power loss), solar array failures
• Un-commanded change in system states (phantom commands)
• Loss of synchronization in timing circuits
• Spurious mode switching, power-on resets, erroneous sensor signals
• Telemetry noise, loss of data
Electromagnetic interference (EMI)
• EMI noise levels in receiver band exceeding receiver sensitivity
• Communications issues due to excess noise
• Phantom commands, signals
Material damage
• ESD damage to mission critical components including
– Thermal control coatings
– Insulating re-entry thermal protection system materials
– Optical materials (dielectric coatings, mirror surfaces)
• Photo-ionized outgassing materials deposited as surface contaminants
Other
• Parasitic currents and solar array power loss (LEO)
• Compromised science instrument function
– Ion spectrum modified by “Ion line” charging signature
– Photoelectron contamination in electron spectrum
– Thermal electron (ion) population cannot reach detectors when spacecraft charged negative (positive) 12
Cable Damage
• Signal, power cable insulation is susceptible to ESD damage
• Exposed biased wires can short to grounded materials
Kawakita et al., 2005
ADEOS-II Failure Investigation
Leung et al., 2010
Damage cable insulation test
Exposed biased conductor
shorts
Leung et al., 2010
13
Solar Array Damage
• Exposed high voltage interconnects interact with plasma environment resulting in current collection
• Energetic electrons charge coverglass, grout, and other dielectric materials with possibility of arcing
• Arc debris can contaminate solar cell and reduce solar illumination
• ESD can lead to sustained arcing, an event that couples solar array current into arc: potential catastrophic event ESA EURECA solar array sustained
arc damage (ESA)
Levy et al., 2001Anti-reflection coating degradation
Gerhard et al., 2001
14
Lab Testing Coverglass Damage
• AFRL lab tests by Ferguson et al. 2016 of GPS-like solar arrays
• Test exposures– 90 keV electrons
– Beam current 0.3x10-12 A/cm2
• 333 arcs observed in 17.5 hours
• Lichtenberg patterns in damaged coverglass material near cell edge
• 100 micron scale
15
Mylar Sheet Arc Damage
• Balmain, 1987 lab tests• Mylar sheet over conducting substrate• 25 keV electrons• Lichtenberg discharge patterns observed
during electron exposure with damage to Kapton
16
ISS Thermal Control Coating
• NASA MSFC arcing test of ISS thermal control coatings used on US sector modules
• 1.3 m chromic acid anodized aluminum
• Coupons biased to voltages of 50 V to 100 V in vacuum with exposure to Argon ions to simulate LEO solar array charging
• 100 V / 1.3x10-6 m ~ 7.7x107 V/m
centimeters
T. Schneider/NASA/MSFC
V
+ + +
+
+
+
+
+
+
+
1.3 m 17
Sensors
• Current work on characterizing charging relies primarily on in-situ sensors that provide information on spacecraft potential and electron, ion environment– No information on discharges or damage
• Flight experiments on surface and internal charging have been flown but only on an occasional basis
• In-space inspection of spacecraft offers the opportunity to detect ESD effects– Best for surfaces exposed to space (surface charging)
– Damage inside of spacecraft due to internal charging will be difficult to evaluate using remote sensing
– Optical sensors can provide information on location and magnitude of damage
– RF sensors can provide information on discharge rates for actively pulsing dielectrics
18
• Low energy background ions accelerated by spacecraft potential show up as sharp “line” of high ion flux in single channel
E = E0 + q
• Assume initial energy E0 ~ 0 with single charge ions (O+, H+) and read potential (volts) directly from ion line energy (eV)
• Accuracy of potential measurement set by energy width and separation of the energy channels used to infer the potential
“Ion Line” Charging Signature, s/c < 0
-646 volts
19
Langmuir Probe
• Current probe techniques have been widely used for many years to measure spacecraft potentials
• Technique is based on measuring current collected by probe as a function of the probe voltage
HH MM SS.MSEC Ni Te Vflt Vsp<------ GPS ------> (m-3) (K) (volt) (volt)
--------------------------------------------------------------------02 20 00.668 7.06e+10 1.95e+03 6.39 7.1002 20 01.668 7.08e+10 1.46e+03 6.39 6.9102 20 02.672 7.30e+10 1.51e+03 6.37 6.9102 20 03.672 7.03e+10 1.24e+03 6.39 6.8102 20 04.672 7.20e+10 1.76e+03 6.32 6.9802 20 05.672 6.92e+10 1.68e+03 6.37 6.9802 20 06.676 7.24e+10 1.61e+03 6.39 6.9302 20 08.676 7.28e+10 1.55e+03 6.32 6.8602 20 09.680 7.20e+10 1.45e+03 6.42 6.9102 20 10.680 7.26e+10 1.56e+03 6.29 6.8802 20 11.680 7.07e+10 1.74e+03 6.37 7.03
--------------------------------------------------------------------
[from Merlino, 2007]
Ii(V)ion
collection
Ie(V)
electron collectionelectron
retardation
Vfloat Vspacewhere 𝐼𝑥,𝑠 = 0.25𝑒𝑛𝑥𝑣𝑥,𝑡ℎ𝐴𝑝𝑟𝑜𝑏𝑒 for x=I,e
Floating Potential Measurement Unit
𝐼𝑖 𝑉𝐵 = ൞−𝐼𝑖𝑠𝑒𝑥𝑝
𝑒 𝑉𝑃 − 𝑉𝐵𝑘𝑇𝑖
, 𝑉𝐵≥ 𝑉𝑃
−𝐼𝑖𝑠, 𝑉𝐵< 𝑉𝑃
𝐼𝑒 𝑉𝐵 = ൞𝐼𝑒𝑠𝑒𝑥𝑝
−𝑒 𝑉𝑃 − 𝑉𝐵𝑘𝑇𝑒
, 𝑉𝐵≤ 𝑉𝑃
𝐼𝑒𝑠, 𝑉𝐵> 𝑉𝑃
0.8 m
20
Charging vs MMOD
• Ability to discriminate between charging (ESD) and MMOD damage is required to correctly attribute results from in-space inspection to damage mechanism
• An issue is that MMOD and charging failures may not be independent
– MMOD strikes can initiate an ESD event from charged materials
– Some thought that hypervelocity impacts can be responsible for RF noise due to ESD and electromagnetic pulse generated by impacts (Close et al., 2013; Garrett and Close, 2013)
• Same optical sensors can be used to inspect for surface charging and MMOD damage
STS-7 window pit
NASA ODPO
Arc pit damage
MMOD
Kawakita et al., 2005
Levy et al., 2001
21
ISS Solar Array
Orbital Debris Quarterly, vol 18, issue 4, 2014 https://www.orbitaldebris.jsc.nasa.gov/quarterly-news/pdfs/odqnv18i4.pdf
ISS 3A solar array wing, MMOD damage to bypass diode results in overheating of cell
22
In-Space Inspection Opportunities
• Anomaly and failure investigations
– Characterize location and magnitude of damage
– Unambiguously determine if damage on solar arrays is the cause of power loss or failure
– Provide information useful for development of better designs that mitigate charging
– Significant inventory of historical spacecraft with failures attributed to charging, particularly in GEO
• Operations
– Satellite servicing missions could benefit from use of sensors to detect differential charging between client and servicing spacecraft
– Direct drive electric propulsion systems using high voltage (100’s V) solar arrays could benefit from on-orbit evaluations of solar array design
– Inspection of operational LEO, GEO spacecraft for ESD damage to evaluate charging designs
23
Questions?
24
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