30-cm mercury ion thruster performance with a 1 kW capacitor-diode voltage multiplier beam supply

Abstract

A 1 kW Solar Array and Capacitor-Diode Voltage Multiplier Converter (S/A-CDVM) has been successfully integrated with a 30 cmdiameter mercury ion thruster system to provide ion beam power. Measurements were made to compare steady state and transient response ,r, performance of a conventional bridge converter with the S/A-CDVM con-°; venter used for the ion beam supply. The ability to recover from screen to accelerator arcs and promptly re-establish stable thruster m performance was demonstrated. Solar array transient response to thruster arcing was measured. INTRODUCTIONIn space flight applications of primary electric propulsion the thrust system mass and efficiency are important considerations. For some planetary missions thruster system mass and efficiency are critical considerations in determining the feasiability for accomplishing the mission. 'the power processors constitute a major portion of the thrust system mass. Its efficiency determines the total solar array power requirement and thermal control mass required. These factors have driven the technology program for lightweight, efficient power processors.A power processor which incorporates the capacitor-diode voltage multiplier (CDVM) concept (refs. 1 to 4) holds promise of fulfilling this need. A CDVM transfers energy from its source to the load by capacitance rather than magnetic coupling, and consequently does not require the weight intensive l ower transformer (and some of the inductors) used in series resonant and conventional converters. The CDVM converter allows the use of a higher switching frequency which reduces the circuit capacitance and filter inductance. By eliminating the transformer, +sing a high switching frequency, and incorporating high energy density capacitors, he mass and dissapative losses of the CDVM converter are significantly reduced over present day power processors. The efficiency of this type of converter has been demonstrated to be typically 95 percent (rcf. 4).The results of a recent study (ref. 5) of a space mission requiring an extended performance electric propulsion system (Comet Malley rendezvous) show that an 11 percent savings in thrust system STAR Category 20 AIAA Paper No. 78-686