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This NASA Technical Standard provides requirements relative to various plasma interactions that can result when a high-voltage system is operated in the Earth's ionosphere and standard practices to eliminate or mitigate such reactions.

Purpose

The purpose of this NASA Technical Standard is to provide a design standard for spacecraft electrical power systems using voltages greater than 55 volts that operate in the low Earth orbit (LEO) plasma environment encountered in altitudes up to 2000 kilometers (km) and latitudes between -50 and +50 degrees. Such power systems, particularly solar arrays, are the proximate cause of spacecraft charging in LEO; and these systems can interact with this environment in a number of ways that are potentially destructive to themselves as well as to the platform or vehicle that has deployed them.

High-voltage systems are used in space for two primary reasons. The first reason is to save launch weight. For the same power level, higher voltages enable use of smaller diameter wires (lighter cabling). This is true because P = IV, and V = IR, so P = I2R (where P is power, I is current, R is resistance, and V is voltage). If I is decreased by use of higher V, then smaller wires can be used with no increase in power loss due to cabling. Of course, if one uses the same cable mass, higher voltages will enable higher efficiencies, since less power will be lost to resistance in the cables. For very large power systems, the decrease in cable mass can be substantial.

The second reason to use a high voltage power system is that some spacecraft functions require them. For example, electric propulsion uses voltages from about 300 V (Hall thrusters) to about 1000 V (ion thrusters). A low-voltage power system would require conversion of substantial power to high voltages 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. If the highvoltage function is electric propulsion, such a system is called a direct-drive electric propulsion system.

Because of these and other reasons, spacecraft designers and manufacturers are increasingly employing high voltage power systems, but with the advantages comes a higher risk of spacecraft charging. The presence of high voltage solar arrays and exposed electrical conductors that carry high voltages can directly exacerbate the spacecraft charging process in LEO, potentially resulting in undesirable electrical arcing, power drain and disruptions, and contamination of spacecraft surfaces and coatings all of which contribute to the damage and loss of spacecraft coatings. Thus, system designers need a standard to show them how to

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mitigate the spacecraft charging effects of using high voltages in LEO. In addition to system designers, this document is useful to space mission personnel including project managers, solar array designers, and system engineers.

This document is intended as a standard for design applications and can be used as a requirements specification instrument.

Applicability

This NASA Technical Standard is applicable to spacecraft electrical power systems using voltages greater than 55 volts that operate in the LEO plasma environment encountered in altitudes up to 2000 kilometers (km) and latitudes between -50 and +50 degrees. Specifically excluded are spacecraft that encounter GEO charging conditions that do not (often) encounter energetic electrons within the auroral ovals, 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 and internal charging will be a concern. However, it is not in the purview of this document to deal with those two topics. For direction on designing spacecraft to survive the conditions in GEO and in the Van Allen belts, one should reference NASA-HDBK- 4002A, Mitigating In-Space Charging Effects-A Guideline.

Some of the design standards for LEO are at variance with good design practice for GEO spacecraft. For example, in GEO one would use materials on the external surface of the spacecraft with low electrical resistance that are all bonded together. This prevents external charging and the potential for electrostatic discharge. In LEO, if you make external materials conductive, then more current is collected from the plasma increasing the parasitic currents in the system and changing the system floating potential. If the spacecraft will fly in both LEO and GEO conditions, be careful to use design solutions that are applicable in both environmental regimes (see NASA-HDBK-4002A).

This NASA Technical Standard is approved for use by NASA Headquarters and NASA Centers and Facilities and may be cited in contract, program, and other Agency documents as a technical requirement. It may also apply to the Jet Propulsion Laboratory and other contractors only to the extent specified or referenced in applicable contracts.

Verifiable requirement statements are numbered and indicated by the word "shall" beginning in section 4; this NASA Technical Standard contains 14 requirements. Explanatory or guidance text is indicated in italics beginning in section 4. To facilitate requirements selection and verification by NASA programs and projects, a Requirements Compliance Matrix is provided in Appendix A.