Spacecraft exploring solar and planetary plasma effects typically carry a sensitive electric-field measurement instrument operating in the tens of Hz to tens of MHz frequency range. These instruments are subject to unique, non-intuitive interference mechanisms driven by the interaction of direct energy transfer electrical power systems with the surrounding plasma. These mechanisms are not addressed by typical spacecraft EMI control programs based on MIL-STD, Aerospace TOR, or AIAA spacecraft EMI requirements.A direct energy transfer spacecraft power system controls electrical power bus voltage and battery load by switching individual solar array circuits on and off of the system power bus based on battery state of charge and spacecraft loads. This arrangement results in a direct electrical connection between solar array strings and the spacecraft power bus, providing a noise propagation path to the solar arrays.The change in solar array string voltage when switched between loaded and unloaded conditions also changes the net potential of the spacecraft surface with respect to the surrounding plasma, resulting in a "bounce" in the voltage relationship between the spacecraft and the surrounding plasma, which is detected by the e-field sensors as noise.This paper describes these phenomena and the techniques used to assess potential effects on the plasma instrument carried on the Juno spacecraft; summarizes the design approach applied to control these effects; and reviews the methods used to verify the effectiveness of the control approach.

No abstract

BEFORE COMPLETING FORM 1. REPORT NUMBER 3GvACKyN~.3 RECIPIENT'S CATAL.OG NUMBER 8 "0 TABLE OF CONTENTS (Cont'd) Section Page Ii. 6 7 TECHNOLOGY FORECAST 7.1 Introduction 7.2 IC Technologies 64 7.3 Electromagnetic Environment 64 7.4 Packaging and Interconnection 64 " 7.5 IC Complexity and Density 65 7.6 Impact on EMC Engineering 65 REFERENCES 66