Spacecraft Charging (also see Spacecraft Charging Flash animation)

Spacecraft charging includes both surface charging and internal dielectric charging. The source and design mitigation techniques are different for each. Spacecraft charging consequences have ranged from intermittent anomalous behavior up to catastrophic satellite failure. The absolute charging of spacecraft surfaces (relative to ambient plasma) is not generally detrimental; rather it is the possible discharge effects which can disrupt satellite operations. Most of the undesired effects of both charging types are due to the discharge arcing, and include physical materials damage and electromagnetic interference (EMI) generation (and resultant transient pulses). Generally, the potential for surface charging problems has been estimated using computer simulation (see Fig. 1) whereas dielectric materials susceptible to internal charging can be tested using electron beam exposure.

nascapfuse

Figure 1: NASA Charging Analysis Program (NASCAP) representation of FUSE satellite
(from EnviroNET at Goddard Space Flight Center).

As a historical perspective, spacecraft charging became an issue after several incidents of anomalous behavior on satellites during the early 1970s, and the complete loss of the DSCS 9431 satellite. As a result, a large program was jointly conducted by NASA and the U.S. Air Force to investigate the problem in the late 1970s. Various missions have been conducted to study spacecraft charging including (a) the Spacecraft Charging AT High Altitudes (SCATHA) satellite that was launched in January 1979 to study spacecraft charging in near geosynchronous orbit, and (b) the Combined Release and Radiation Effects Spacecraft (CRRES) that was launched in July 1990 into a geosynchronous transfer orbit.

The two categories of spacecraft charging addressed below are

  1. surface charging (also includes differential charging), and
  2. internal dielectric (bulk, buried, deep, or thick) charging.

The spacecraft charging is driven by charged particle motion in and transfer from the nearby plasma. These plasma energies are from eV to keV levels, as compared to the MeV particles typical of ionizing radiation. Differential surface charging has been reported to correlate best with the intensity of electrons with E <= 50 keV.[1,2] Electrons with energy > 50 keV can penetrate spacecraft surface metallization to cause internal discharge.[3]


Surface Charging

Surface charging is created from low-energy plasma and photoelectric currents (see Fig. 2). The midnight to dawn sector is a favored region for surface charging-induced anomalies. Typically, differential charging has occurred after geomagnetic substorms, which result in the injection of keV electrons into the magnetosphere. While in eclipse, the spacecraft may negatively charge to tens of kilovolts. A potential sufficient for discharge is easily created when the satellite emerges into sunlight, which results in positive surface charge due to photoelectron emission. Differential charging can also be caused by satellite self-shadowing. The basic solution to differential charging problems is to provide a common ground for the spacecraft surface (including internal structures). Spacecraft in geosynchronous orbit are more likely to undergo differential charging. However, use of a high-voltage power system in a low earth orbit satellite can increase the adverse environmental effects.

charge

Figure 2: Charge control mechanism for a satellite
developed by the Austrian Research Centre Seibersdorf.

The spacecraft surface potential is a function of the net current flow to/from the spacecraft surface. Figures 2 and 3 illustrate the spacecraft charging process. These currents are from solar photon-induced photoelectrons leaving the surface, plasma electrons and ions impinging on the surface, and charged particles emitted from the vehicle (e.g., from active ion emission). A balance equation for current density can be written as:

Jelec + Jion + Jpe + Jsec + Jback + Jart = 0 (1)

where the currents Jelec and Jion are from external plasma electrons and ions, respectively, Jpe is the net photoelectron current, Jsec is the net current due to secondary electrons (few eV) generated by energetic primaries (electrons and ions) at the satellite surface, Jback is the backscattered electron current from electrons reflected back from the surface with some energy loss, and Jart is a possible artificial current.

A spacecraft placed in the plasma will assume a floating potential different from the plasma itself.[4] Figure 3 illustrates surface charging in both sunlight and shadow. In darkness, a spacecraft surface will tend to charge negatively from the ambient plasma electrons. The plasma is basically neutral, having equal numbers of electrons and ions at equal energies; however, the electrons are much lighter particles, and therefore, move at higher velocities (i.e., EKmv²). Hence, the negative electron current to the spacecraft surface is greater than the positive ion current.[5]

SurfaceCharging

(a) Surface in Shadow: the current balance requires equality between the flow of the plasma ions and that of the plasma electrons impinging on the surface. (b) Surface in Sunlight: equilibrium is achieved when the flow of escaping photoelectrons is equal to the difference between the incoming flows of plasma electrons and ions.


Figure 3: Qualitative illustration of surface charging by plasma. The arrow widths are proportional to the flux of each particle species; the equilibrium potential is reached when the sum of the currents collected and emitted by a surface element is zero. [6]

A rough estimate of the current to the spacecraft due to the ambient plasma can be made. If the plasma is assumed at a single temperature and density, the Maxwell-Boltzmann distribution may be used to represent the ambient conditions.[2] In this case the average velocity is [3]

vavg = [(2 k Ti)/(pi mi)]1/2 (2)

The current flux from either electrons or ions can then be computed from:

Ji = qi Ni vavg / 2 (3)

where qi is the charge, and Ni is the number density. Garrett continues such an analytic derivation to find that in eclipse, the potential between the satellite and space is [2]

V -TE (4)

where TE is the ambient electron temperature (in eV).

Spacecraft charging is a function of the space environment characteristics, including sunlight/eclipse, solar activity, geomagnetic activity, electron flux magnitude and spectrum. These effects can be (dis)advantageous, for instance, sunlight exposure provides photoemission, that can act as a charge drain to neutralize the surface potential, or that can act as a discharge trigger upon emergence from eclipse/shadow. Table I is a reproduction of extreme satellite potentials in different plasma environments provided by Grard et al. McPherson and Schober indicate for sun exposure that at lower altitudes (< 2 RE) the flux of plasma electrons to the satellite is greater than the photoelectric flux, so the satellite becomes negatively charged; whereas, for higher altitudes (> 3 RE) the photoelectric flux dominates and the spacecraft becomes positively charged.[7]

Table I: Extreme Potentials for Satellites in Different Plasma Environments [6]

Environment Satellite in eclipse, or insulated surface in shadow Sunlit satellite with conductive surfaces
Plasmasphere (out to ~ 5 RE) -2 V +2 V
Dayside magnetosphere (~ 5 to ~ 10 RE) -5 to -100? V +10 V
Nightside magnetosphere (~ 5 to ~ 15 RE) -20,000 V (during magnetospheric storms) +30 V
Dayside magnetosheath (~ 10 to ~ 15 RE) -200 V +5 V
Solar wind -20 V +10 V

For surface charging, the surface typically charges to the mean energy (in eV) of the dominant surface current (see Eq. 4). At geosynchronous orbit (GEO), surfaces exposed to sunlight charge to 2-3 volts (positive) due to the photoelectron current emitted from the surface; during eclipses, a negative surface potential is observed.[2] Photoemission from the extreme ultraviolet wavelength range (< 2000 Å) is the most important since in that region many materials have rather large photoelectric yields and the solar spectrum also has significant energy there.[8] In eclipse the spacecraft roughly charges to a negative potential equivalent to the electron temperature (i.e., kT). Shadowed dielectric or isolated surfaces, the potential may charge to 1 to 10 kV (negative) from local electrons. Surface discharges are primarily caused by low energy electrons (up to a few tens of keV). In low Earth orbit (LEO), the thermal electron currents are the largest and satellites tend to be slightly negative.[3] Whipple summarizes equilibrium potentials at increasing altitude:[8]

Equilibrium Potentials at Increasing Altitude
Ionosphere: a few tenths of a volt negative
Magnetosphere: normally, a few volts positive; in eclipse, may become highly negative
Solar Wind: a few volts positive
Interstellar Space: a few volts positive or negative

Absolute v. Differential Charging

Absolute charging occurs when the satellite potential relative to the ambient plasma is changed uniformly. Differential charging occurs when parts of the spacecraft are charged to different potential relative to one another. Absolute charging is on the order of microseconds. Differential charging typically occurs over seconds to minutes because of capacitance considerations. The absolute charging of spacecraft surfaces (relative to ambient plasma) is not generally detrimental; rather it is the possible discharge effects from differential charging (see Fig. 4) which can disrupt satellite operations. Most of the undesired effects of charging are due to the discharge arcing, and include physical materials damage and EMI generation (and resultant transient pulses). Arc-discharges occur when the electric fields created by differential charging exceed breakdown potentials. The arcs are rapid (~ nanosec) and rearrange charge distribution by punch-through (internal dielectric breakdown), and by flashover between surfaces or between surfaces and space. Shaw et al. demonstrated that the average satellite discharge rate increases with increasing geomagnetic activity.[9] Discharge consequences include noise in both data and wiring. Arcing, although seen as the primary mechanism by which spacecraft charging disturbs mission activities, is not the only consequence of spacecraft charging. Other negative effects include sputtering and attraction of chemically active species.

DifferentialCharging

Figure 4: Differential Charging due to Self-shadowing.[10]


Internal Dielectric Charging

Internal (deep or bulk) dielectric charging is caused by high-energy electrons penetrating dielectric materials (e.g., printed circuit boards). These high-energy electrons are more likely found trapped within the earth's Van Allen radiation belts. Normally, a fluence of 1010 - 1011 electrons/cm² (over a period relative to the dielectric leakage rate) will build-up a sufficient charge for arcing. Some researchers have indicated a higher likelihood of deep dielectric charging-induced anomalies than those from surface charging and single-event upset. Leaky dielectrics, proper grounding, and shielding can be used to reduce the possibility of internal charging. In addition, EMI-susceptibility reduction techniques can be employed to mitigate the effects of arcing.

The fact that the explanation given here for internal charging is shorter than that for surface (differential) charging should not be indicative of the relative importance of the two. In fact, just the opposite, internal discharge is more damaging since it occurs within dielectric materials and well-insulated conductors, which are in close proximity to sensitive electronic circuitry.[5] Gussenhoven et al. state that based on CRRES data obtained at GEO, most environmentally induced spacecraft anomalies result from deep dielectric charging and the resulting discharge pulses and not from surface insulator charging or single-event upsets.[11] In addition, the mechanisms for internal charging are more straightforward, i.e., high-energy electrons penetrate internal dielectric materials, and if charge buildup occurs too rapidly, then an arc discharge ensues.

High-energy (E > 100 keV) electrons may penetrate into the satellite, and establish negative potential on isolated parts such as dielectric materials and floating conductors. The electrons may become trapped (buried) in dielectric materials. Kapton™ and Teflon™ are dielectric materials used extensively on spacecraft (as thermal blankets) which poorly distribute electric charge. The internal electric field will build up if the charge leakage rate is less than the charge collection rate. A fluence of 1010 - 1011 electrons/cm² (over a time period relative to the dielectric leakage rate) will build-up a sufficient charge for arcing.[12] The resulting arcing will appear as a pulse on the cabling and circuit board. Pulse widths are usually in the tens of nanoseconds. Also, printed circuit boards with islands of metallization will charge up like a capacitor. If a sufficient potential is reached, arcing may result in upset or burnout of nearby semiconductor devices.

The internal charging can affect insulators such as cable wrap, wire insulation, circuit boards, electrical connectors, feed throughs, etc. The likelihood of discharge is a function of both the voltage potential and the electric field. Some simple relationships between these factors can be established. The current density (J) to the dielectric is obtained from the electron flux (phi):

J [A/cm²] = phi [e/cm²·sec] q [C/e] (5)

where q=1.6022x10-19 C/e. If a simple capacitor plate behavior is assumed, the electric field buildup may be determined from the conductivity of the material:

E(t) = V(t)/d = J/s [1 - exp(-t/t)] (6)

where s is dielectric conductivity, t is the relaxation time, and d is the thickness of the material. This electric field (or the maximum field of J/s) may be compared to the breakdown strength of the dielectric to determine whether or not arcing will occur.


Orbital Considerations

Two important orbits for which more charging data exists are the low earth orbit (LEO) and geosynchronous orbit (GEO). LEO is around 100 to 2000 km altitude; GEO is about 36,000 km (or at 6.6 RE). Important orbital variations in the plasma environment are shown in Table II and Figure 5.

Table II: Typical Plasma Parameters for LEO and GEO [13]

Plasma Parameter GEO LEO
Density (m-3) 106 1011
Temperature (eV) 103 0.2
Electron Thermal Current (A/m²) 10-6 10-3
Ram Ion Current (A/m²) 5x10-10 10-4

At LEO spacecraft experience low-energy, high-density electrons; whereas GEO spacecraft encounter high-energy, low-density electrons. Energy level affects voltage potential (see Eq. 4) whereas density determines charging current density (see Eq. 5). A satellite in a low altitude, high inclination (polar) orbit will pass through auroral zones, which have low-density, high-energy electrons and ions.

SpacecraftChargingOrbital
Figure 5: Properties of the natural space plasma. [10]

LEO Wake Effects

For some altitudes, the relative velocity between the satellite and the ambient plasma can become significant enough to create a plasma wake. Below 800 km (i.e., LEO) wake effects may be significant for ions such that an asymmetry exists between the ion flux to the leading edge and rear surface of the satellite.[14] Laframboise and Luo discuss the wake-induced barrier effect mechanism in detail for a dielectric spacecraft (e.g., the Shuttle Orbiter) in polar orbit.[15]

In LEO, spacecraft move through a dense, low-energy plasma (see Table II). LEO spacecraft are negatively charged because their orbital velocity (8 km/s) is greater than the ion thermal velocity (1 km/s) but slower than the electron thermal velocity (200 km/s).[16] Hence, a wake effect in which ions can only impact ram surfaces while electrons are capable of impinging on all surfaces. The wake effect results in regions of differential charging.


References

  1. S.E. DeForest, "Spacecraft charging at synchronous orbit," Journal of Geophysical Research, vol. 77, no. 4, pp. 651-659, Feb. 1972.
  2. H.B. Garrett, "The charging of spacecraft surfaces," Reviews of Geophysics and Space Physics, vol. 19, no. 4, pp. 577-616, Nov. 1981.
  3. P.A. Robinson, Jr., P. Coakley, "Spacecraft charging: Progress in the study of dielectrics and plasmas," IEEE Trans. on Electrical Insulation, vol. 27, no. 5, pp. 944-960, Oct. 1992.
  4. A.R. Martin, "A review of spacecraft/plasma interactions and effects on space systems," Journal of The British Interplanetary Society, vol. 47, no. 4, pp. 134-142, April 1994.
  5. R.D. Leach, M.B. Alexander, Failures and Anomalies Attributed to Spacecraft Charging, NASA RP-1375, Marshall Space Flight Center, AL, Aug. 1995.
  6. R. Grard, K. Knott, Pedersen, "Spacecraft charging effects," Space Science Reviews, vol. 34, pp. 289-304, 1983.
  7. D.A. McPherson, W.R. Schober, "Spacecraft charging at high altitudes: The SCATHA satellite program," from the AIAA Symposium on "Spacecraft Charging by Magnetospheric Plasmas" held in Washington, D.C., June 1975, Progress in Astronautics and Aeronautics, vol. 47, pp. 15-30, MIT Press, 1976.
  8. E.C. Whipple, "Potentials of surfaces in space," Reports on Progress in Physics, vol. 44, pp. 1197-1250, 1981.
  9. R.R. Shae, J.E. Nanevicz, R.C. Adamo, "Observations of electrical discharges caused by differential satellite-charging," from the AIAA Symposium on "Spacecraft Charging by Magnetospheric Plasmas" held in Washington, D.C., June 1975, Progress in Astronautics and Aeronautics, vol. 47, pp. 61-76, MIT Press, 1976.
  10. J.L. Herr, M.B. McCollum, Spacecraft Environments Interactions: Protecting Against the Effects of Spacecraft Charging, NASA RP-1354, Marshall Space Flight Center, AL, Nov. 1994.
  11. M.S. Gussenhoven, E.G. Mullen, D.H. Brautigam, "Improved understanding of the Earth's radiation belts from the CRRES satellite," IEEE Trans. on Nuclear Science, vol. 43, no. 2, pp. 353-368, April 1996.
  12. P. Leung, A.C. Whittlesey, H.B. Garrett, P.A. Robinson, Jr., "Environment-induced electrostatic discharges as the cause of Voyager 1 power-on resets," Journal of Spacecraft and Rockets, vol. 23, no. 3, pp. 323-330, May-June 1986.
  13. M.J. Mandell, I. Katz, D.L. Cooke, "Potentials on large spacecraft in LEO," IEEE Trans. on Nuclear Science, vol. NS-29, no. 6, pp. 1584-1588, Dec. 1982.
  14. H.B. Garrett, "Spacecraft charging: A review," in Space Systems and Their Interactions with Earth's Space Environment, H.B. Garrett, C.P. Pike, eds., Progress in Astronautics and Aeronautics, vol. 71, AIAA, pp. 167-226, 1980.
  15. J.G. Laframboise, J. Luo, "High-voltage polar-orbit and beam-induced charging of a dielectric spacecraft: A wake-induced barrier effect mechanism," Journal of Geophysical Research, vol. 94, no. A7, pp. 9033-9048, July 1989.
  16. B.F. James, O.W. Norton, M.B. Alexander, The Natural Space Environment: Effects on Spacecraft, NASA RP-1350, Marshall Space Flight Center, AL, Nov. 1994.


Last updated: January 19, 2006
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