Spacecraft Coating-Induced Charging: Materials and Modeling Study of Environmental Extremes

Abstract

As mankind reaches to explore extreme environments in space, the application of ceramics surface coatings is increasing. The 2005 mission concept for Solar Probe used a unique design to achieve the necessary thermal control for a very close approach to the solar corona, including the use of a highly refractory, electrically insulating ceramic coating over a carbon-carbon composite heat shield. The proposed trajectory takes the spacecraft from a Jovian fly-by to within 4 solar radii of the Sun, spanning 5 orders of magnitude in solar radiation and solar wind plasma density as well as spacecraft temperatures from <100 K to >2000 K. Using the NASCAP-2K charging modeling program, the degree of charging expected for this spacecraft design has been calculated for this range of radiation environments. New measurements of the electron emission and estimates of related properties of the candidate materials-Al 2 O 3 , pyrolytic born nitride and barium zirconium phosphate-are presented. Absolute and differential surface charging are found to depend strongly on temperature through increased conductivity at higher temperatures and on radiation flux through enhanced charge accumulation and radiation induced conductivity. As the spacecraft approaches the Sun, the competition between increased charge dissipation at higher temperatures and increased charge accumulation at higher fluxes leads to a maximum in differential charging between 0.4 AU and 2 AU. While the spacecraft charging behavior of these materials is found to be significant, it is not severe enough to endanger the mission, and a number of options exist to mitigate the degree of charging. Among the ceramics considered, the use of Al 2 O 3 coatings is found to minimize both absolute and differential spacecraft charging. [Type text] 2 E max = energy at maximum secondary emission δ max occurs el inc E = incident electron energy BSE emit J = emitted current due to backscattered electron emission ion emit J , photo emit J = emitted current due to ion-induced and photon-induced secondary electron emission SE emit J = emitted current due to electron-induced secondary electron emission el inc J , ions inc J = incident electron and ion currents el inc J is the incident electron current, ions inc J is the incident ion current, SE emit J is the current out due to electroninduced secondary electron emission, BSE emit J is the current out due to backscattered electron emission, ion emit J is the current out due to ion-induced secondary electron emission, and photo emit el inc net J J J J J J J + + + − + =(1) Note that in our convention electron currents are negative due to the sign of the charge carrier.At equilibrium, J net = 0, such that the current into the spacecraft (in the form of incident ions and electrons) equals the current out of the spacecraft (due to photoemission, secondary electron emission, and backscattered electrons) such that a potential develops relative to the surrounding plasma "ground." Absolute charging develops when the local "ground" of the entire spacecr...

Parker Solar Probe FIELDS Instrument Charging in the Near Sun Environment: Part 2: Comparison of In‐Flight Data and Modeling Results

“…The FIELDS shields protect the stub of the antenna as it connects to the hinge and from there to the rest of the FIELDS instrument which is shielded from the Sun by the TPS. Previous PSP models did not include the entirety of the FIELDS instrument nor geometry and did not have the proper material properties to perform a full spacecraft charging simulation (Donegan et al, 2010(Donegan et al, , 2014Ergun et al, 2010;Guillemant et al, 2012).…”

Parker Solar Probe FIELDS Instrument Charging in the Near Sun Environment: Part 1: Computational Model

“…Previous PSP spacecraft surface charging models involved the PSP spacecraft, including the Thermal Protection System (TPS), the spacecraft radiators, and the bus. However, they did not include the FIELDS antennas (Donegan et al, 2010(Donegan et al, , 2014Ergun et al, 2010;Guillemant et al, 2012) because the necessary information to model the antennas, such as probe surface properties were not available, and the final geometry of these thin (0.0031 m diameter) 2 m long probes were unknown at the time of their publishing.…”