Figure 1: NASA Charging Analysis Program (NASCAP) representation of FUSE satellite
(from EnviroNET at Goddard Space Flight Center).
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.
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
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 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.
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., EK=½mv²). Hence, the negative electron current to the spacecraft surface is greater than the positive ion current.[5]
| (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. |
| |
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)/( 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 |
Figure 4: Differential Charging due to Self-shadowing.[10]
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 (
):
| J [A/cm²] = [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/ [1 - exp(-t/ )]
| (6) |
where is dielectric conductivity,
is the relaxation time, and
d is the thickness of the material. This electric field (or the maximum field
of J/) may be compared to the
breakdown strength of the dielectric to determine whether or not arcing will occur.
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.
Figure 5: Properties of the natural space plasma. [10]
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.
