scope:

This document and information handbook presents an overview of the current understanding of the various plasma interactions that can result when a high-voltage system is operated in the Earth's ionosphere. This document is also applicable to satellites during the lower latitude portion of polar orbits and to GEO satellites during the initial low altitude portion of their geostationary transfer orbit.

It references common design practices that have exacerbated plasma interactions in the past, and recommends standard requirements and practices to eliminate or mitigate such reactions.

Purpose

The purpose of this standard is to provide the requirements for a design standard for high-voltage space power systems (>55 V) that operate in the plasma environment associated with LEO (altitude between 200 km and 1000 km and latitude between -50 degrees and +50 degrees). Such power systems, particularly solar arrays, are the proximate cause of spacecraft charging in LEO. These systems can interact with this environment in a number of ways that are potentially destructive to themselves and to the platform or vehicle that has deployed them.

High-voltage systems are used in space for two reasons. The first is to save launch weight. High-voltage systems are often used as a means to reduce mass and increase efficiencies by reducing power line I2R losses. High-voltage systems are also used in space because some spacecraft functions require high voltages. For example, electric propulsion uses voltages from about 300 V (Hall thrusters) to about 1000 V (ion thrusters). For low-voltage power systems, conversion of substantial power to high voltages is required for these spacecraft functions to operate. The weight of the power conversion systems, power management and distribution (PMAD), can be a substantial fraction of the total power system weight in these cases. It is more efficient, and can save weight, if the high-voltage functions can be directly powered from a high-voltage solar array, for instance. If the high-voltage function is electric propulsion, then we call such a system a direct-drive electric propulsion system.

These two reasons and others are causing spacecraft designers and manufacturers to use high voltages more and more. However, doing so entails risk; in particular, spacecraft charging in LEO, in contrast to that in GEO, is caused by exposed high voltages and can lead to arcing, power drains, power disruptions, and loss of spacecraft coatings. Thus, system designers need a standard to show them how to mitigate the spacecraft-charging effects of using high voltages in LEO. In addition to system designers, this document should be useful to project managers, solar array designers, system engineers, etc.

This document is intended to provide requirements and associated best practices for design applications and can be used as a requirements-specification instrument.

Applicability

The contents of this document are applicable to high-voltage space power systems that operate in the plasma environment associated with LEO. This standard is intended for space systems that spend the majority of their time at altitudes between 200 km and 1000 km (usually known as LEO applications) and at latitudes between about -50 degrees and +50 degrees-that is, space systems that do not encounter GEO charging conditions, that do not (often) encounter the auroral ovals of electron streams, and that do not fly through the Van Allen belts. For the extreme radiation protection that is necessary for orbits in the Van Allen belts, exterior spacecraft charging will likely be a secondary concern. However, internal charging will be very important. It is not in the purview of this document to deal with internal charging.

Some of the design standards for LEO are at variance with good design practice for GEO spacecraft. If your spacecraft will fly in both LEO and GEO conditions, then be careful to use design solutions that are applicable in both environmental regimes.