On-Orbit Detection of Spacecraft Charging Effects

Transcription

1 On-Orbit Detection of Spacecraft Charging Effects Joseph I. Minow NASA, Marshall Space Flight Center In-Space Inspection Workshop 2017 NASA JSC, Houston, Texas 30 January 2 February 2017 ISS image: 7 March

2 Introduction Space Environment Impacts on Space Systems [Koons et al., 2000]* Anomaly Diagnosis Number % ESD-Internal, surface and uncategorized SEU (GCR, SPE, radiation belt) Radiation dose Meteoroids, orbital debris Atomic oxygen Atmospheric drag Other Total % *Sources: Aerospace Corporation study using data from 1) NOAA/NGDC Spacecraft Anomaly Manager 2) NASA Anomaly Reports 3) USAF 55 th Space Weather Squadron anomaly database 4) Individual Program Offices databases 2

3 Introduction Space Environment Impacts on Space Systems [Koons et al., 2000]* 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 and uncategorized SEU (GCR, SPE, radiation belt) Radiation dose Meteoroids, orbital debris Atomic oxygen Atmospheric drag Other Total % *Sources: Aerospace Corporation study using data from 1) NOAA/NGDC Spacecraft Anomaly Manager 2) NASA Anomaly Reports 3) USAF 55 th Space Weather Squadron anomaly database 4) Individual Program Offices databases 3

4 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 Jupiter Anom DMSP F LEO Anom SCATHA 1982 GEO Anom Telstar , 1997 GEO Anom/LOM GOES GEO LOM TSS-1R 1996 LEO Failure AUSSAT-A1, -A2, -A GEO Anom TDRS F GEO Anom FLTSATCOM GEO Anom TDRS F-3,F GEO Anom GOES GEO Anom/SF INSAT GEO Anom/LOM Feng Yun 1A 1988 LEO Anom/LOM Tempo GEO LOM MOP-1, GEO Anom PAS GEO LOM GMS GEO Anom Feng Yun 1C 1999 LEO Anom BS-3A 1990 GEO Anom Landsat LEO Anom MARECS A 1991 GEO LOM ADEOS-II 2003 LEO LOM Anik E GEO Anom/LOM TC-1, ~2xGTO, GTO Anom Anik E GEO Anom Galaxy GEO Anom Intelsat GEO Anom Echostar GEO Anom SAMPEX LEO Anom Suomi NPP LEO Anom *Anom=anomaly, SF=system failure, LOM=Loss of mission 4

5 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 Surface charging dq dt d C dt k I k ~ 0 at equilibrium 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 Internal (deep dielectric) charging D E ( ) 2 Electrodynamic potentials F q( E v B) F qe' E E v B E ds m C [c.f., Whipple, 1981; p. 272 Wangness, 1986; p. 210 Jackson, 1975; Maynard, 1998] C C ( E v B) ds Laboratory frame Spacecraft rest frame Forces equal in both frames! ( E v B) ds 5

6 Surface Charging Physics Surface charging is a current balance process to and from spacecraft surfaces as a function of the spacecraft potential dq dt dq dt C I I I dv dt I k I I I I i e c se si k (V) (V) bs,e (V) (V) (V) ph,e dσ dt (V) (V) A k I k incident ions incident electrons backscattered electrons conduction currents secondary electrons due to I e secondary electrons due to I i photoelectrons few ev to ~50 kev (Garrett and Minow, 2004) 6

7 Internal Charging Physics D ρ t J D εe, J R J R ρ σ radiation ε J J C σ dark k κε 0 J R σe dγ α 0.5 α 1.0 dt σ radiation E ~50 kev to few MeV 7

8 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) Image: JHU/APL, NASA Image: NOAA 8

9 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) NOAA SWPC >30 kev >300 kev NOAA SWPC 9 NOAA SWPC

10 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-6 X SCATHA LANL [Olsen, 1983] 10 (Lai et al, 2003)

11 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 = ½CV 2 inches PMMA (acrylic) charged by ~2 to 5 MeV electrons 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 11

12 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

13 Cable Damage Signal, power cable insulation is susceptible to ESD damage Exposed biased wires can short to grounded materials ADEOS-II Failure Investigation Exposed biased conductor shorts Kawakita et al., 2005 Leung et al., 2010 Damage cable insulation test Leung et al.,

14 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) Gerhard et al., 2001 Levy et al., 2001 Anti-reflection coating degradation 14

15 AFRL lab tests by Ferguson et al of GPS-like solar arrays Test exposures 90 kev electrons Beam current 0.3x10-12 A/cm arcs observed in 17.5 hours Lab Testing Coverglass Damage Lichtenberg patterns in damaged coverglass material near cell edge 100 micron scale 15

16 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

17 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 ~ 7.7x10 7 V/m T. Schneider/NASA/MSFC V centimeters 1.3 m 17

18 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

19 Ion Line Charging Signature, s/c < 0 Low energy background ions accelerated by spacecraft potential show up as sharp line of high ion flux in single channel E = E 0 + q Assume initial energy E 0 ~ 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 -646 volts 19

20 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 I i V B I e V B = I isexp e V P V B kt i, V B V P I is, V B < V P = I esexp e V P V B kt e, V B V P I es, V B > V P where I x,s = 0.25en x v x,th A probe for x=i,e [from Merlino, 2007] electron retardation I i (V) ion collection Vfloat I e (V) electron collection HH MM SS.MSEC Ni Te Vflt Vsp < GPS > (m-3) (K) (volt) (volt) e e e e e e e e e e e e e e e e e e e e e e Vspace Floating Potential Measurement Unit 0.8 m 20

21 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 MMOD STS-7 window pit 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) NASA ODPO Same optical sensors can be used to inspect for surface charging and MMOD damage Arc pit damage Kawakita et al., 2005 Levy et al.,

22 ISS Solar Array ISS 3A solar array wing, MMOD damage to bypass diode results in overheating of cell Orbital Debris Quarterly, vol 18, issue 4,

23 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

24 Questions? 24