In 1997, two commercial geostationary satellites experienced a new phenomenon: sustained solar array arcing. Although arcing on solar arrays in space had been expected from ground tests and space flight experiments, it was heretofore unknown that arcs into the space plasma could turn into arcs between adjacent solar array strings at high interstring voltages. Experiments validated the concept, and design changes were made to succeeding satellites that have prevented similar occurrences. Here, the dates and times of 32 sustained arcing string failures from 1997 to 2002 are reported, and NASCAP-2k charging simulations are used to try to determine thresholds necessary for the arcs to occur. Surprisingly, in some cases, conditions for sustained arcs last for a longer time when the environment is relatively benign. A plausible scenario for primary arcs in sustained arcing conditions is presented that is based on charging time histories from NASCAP-2K simulations. The charging scenario involves solar array voltage turn-on after eclipse or upon unshunting the array, and it hypothesizes lower than usual arc voltage thresholds immediately after differential string voltages are activated. Laboratory testing under both geosynchronous-Earth-orbit and low-Earth-orbit conditions is cited to show plausibility of this argument. Improved design and testing rules for spacecraft to avoid sustained arcing with a minimum of overdesign are presented.

Since electrons are the primary charged particles at the geosynchronous Earth orbit (GEO), understanding of their interactions with spacecraft materials, such as polyimide (PI, Kapton-H®), is important. Understanding of the chemical nature of electron damage and its effect on PI’s electrical and optical properties is still limited. Thus, predictive spacecraft models (electrical charging, thermal, etc) are restricted to only pristine material properties. This is a major source of error in spacecraft construction and anomaly resolution, since material properties change after exposure to the space environment. Ground based measurements are critical to understanding the dynamics of spacecraft materials however it will be shown in this work that standard material handling practice and exposure to air are unacceptable for these studies.

This study measures Radiation Induced Conductivity (RIC) of Low Density Polyethylene (LDPE) over temperatures ranging from similar to 110 K to similar to 350 K. RIC occurs when incident ionizing radiation deposits energy and excites electrons into the conduction band of insulators. Conductivity was measured when a voltage was applied across vacuum-baked, thin film LDPE polymer samples in a parallel plate geometry. RIC was calculated as the difference in sample conductivity under no incident radiation and under an incident similar to 4 MeV electron beam at low incident fluxes of 10(-4)-10(-1) Gr/Sec. The steady-state RIC was found to agree well with the standard power law relation, sigma(RIC) = k(RIC) . (D) over dot(Delta) between conductivity, sigma and adsorbed dose rate, (D) over dot. Both the proportionality constant, k(RIC), and the power, Delta, were found to be temperature dependant above similar to 250 K, with behavior consistent with photoconductivity models developed for localized trap states in disordered semiconductors. Below similar to 250 K, kRIC and Delta exhibited little change. The observed difference in temperature dependence might be related to a structural phase transition seen at T-beta similar to 256 K in prior studies of mechanical and thermodynamic properties of LDPE.

[Abstract] Modest changes in spacecraft charging conditions can lead to abrupt changes in the spacecraft equilibrium, from small positive potentials to large negative potentials relative to the space plasma; this phenomenon is referred to as threshold charging. It is well known that temporal changes of the space plasma environment (electron plasma temperature or density) can cause threshold charging. Threshold charging can also result from by temporal changes in the juxtaposition of the spacecraft to the environment, including spacecraft orbit, orientation, and geometry. This study focuses on the effects of possible changes in electron emission properties of representative spacecraft materials. It is found that for electron-induced emission, the possible threshold scenarios are very rich, since this type of electron emission can cause either positive or negative charging. Alternately, modification of photonor ion-induced electron emission is found to induce threshold charging only in certain favorable cases. Changes of emission properties discussed include modifications due to: contamination, degradation and roughening of surfaces and layered materials; biasing and charge accumulation; bandstructure occupation and density of states caused by heat, optical or particle radiation; optical reflectivity and absorptivity; and inaccuracies and errors in measurements and parameterization of materials properties. An established method is used here to quantitatively gauge the relative extent to which these various changes in electron emission alter a spacecraft’s charging behavior and possibly lead to threshold charging. The absolute charging behavior of a hypothetical flat, twodimensional satellite panel of a single material (either polycrystalline conductor Au or the polymeric polyimide KaptonTM H) is modeled as it undergoes modification and concomitant changes in spacecraft charging in three representative geosynchronous orbit environments, from full sunlight to full shade (eclipse) are considered.