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