Jason D. Frieman

Preliminary Assessment of the Role of a Conducting Vacuum Chamber in the Hall Effect Thruster Electrical Circuit

The role of the electrically conductive vacuum chamber wall in the completion of the discharge circuit of a Hall effect thruster (HET) is experimentally investigated. The Aerojet Rocketdyne T-140 laboratory-model HET operating at a discharge voltage of 300 V, discharge current of 5.16 A, and anode flow rate of 5.80 mg/s serves as a representative HET test bed. The nominal facility operating pressure during thruster firings is 4.9 × 10 -6 Torr corrected for xenon. Two 0.91 m x 0.91 m square aluminum plates are placed adjacent to, but electrically isolated from, the walls of the stainless steel vacuum chamber at two locations with respect to the center of the thruster exit plane: 4.3 m axially downstream along thruster centerline and 2.3 m radially outward centered on the exit plane. The plates are configured in three distinct electrical configurations with corresponding measurements: a) electrically grounded plates with measurements of currents to ground, b) electrically isolated plates with measurements of floating voltages, and c) isolated but electrically connected plates with measurements of the current conducted between them. The measurements are all taken simultaneously with the discharge current oscillations of the thruster at a sampling frequency of 100 MHz. Measurements of the current conducted to ground in the electrically grounded configuration reveal that the axial and radial plates collect ion currents that are 13.6% and 10.7% of the discharge current, respectively; the collected current is coupled to the discharge current oscillations but is smaller in magnitude and phase-delayed. In the electrically connected plate configuration, 5.5% of the average discharge current is observed to flow from the axial plate to the radial plate driven by a floating voltage difference of 7.6 V; this current is uncorrelated in time with the discharge current oscillations. These results indicate that the vacuum chamber conducts current and is a recombination site for a significant number of plume ions during HET operation.

Recommended Practice for Flow Control and Measurement in Electric Propulsion Testing

Accurate control and measurement of propellant flow to a thruster is one of the most basic and fundamental requirements for operation of electric propulsion systems, whether they be in the laboratory or on flight spacecraft. Hence, it is important for the electric propulsion community to have a common understanding of typical methods for flow control and measurement. This paper addresses the topic of propellant flow primarily for the gaseous propellant systems that have dominated laboratory research and flight application over the last few decades, although other types of systems are also briefly discussed. Although most flight systems have employed a type of pressure-fed flow restrictor for flow control, both thermal-based and pressure-based mass flow controllers are routinely used in laboratories. Fundamentals and theory of operation of these types of controllers are presented, along with sources of uncertainty associated with their use. Methods of calibration and recommendations for calibration process...

Impact of Cathode Position and Electrical Facility Effects on Hall Effect Thruster Performance and Discharge Current Behavior

The goal of this investigation is to characterize the electrical interaction between a Hall effect thruster (HET) and an electrically-conductive vacuum chamber. In order to control the strength of this electrical facility interaction, the cathode radial position with respect to the thruster centerline is varied. The Aerojet-Rocketdyne T-140 laboratory-model HET, operating at 300 V with a discharge current of 10.16 A and a mass flow rate of 11.61 mg/s of xenon, serves as the representative HET test bed. The chamber pressure during operation is 7.3 x 10 Torr corrected for xenon. Two 0.91-m x 0.91-m square aluminum plates are placed adjacent to, but electrically isolated from, the walls of the electricallyconductive vacuum chamber at two locations: 1) 2.3 m radially outward from thruster centerline centered along the exit plane and 2) 4.3 m axially downstream from the thruster exit plane. The plates are configured in three distinct electrical configurations (grounded, isolated, and connected to each other). The plates serve as a diagnostic that can be used to characterize the chamber-thruster electrical interaction. The HET body was configured in two configurations: electrically floating and electrically grounded. At each plate configuration and thruster body configuration, the radial separation distance between the cathode and the thruster is varied from 18.1 cm to 128.8 cm away from thruster centerline. At each cathode position, the current-to-ground, the floating voltage, and the current conducted between the plates are measured temporally. At each experimental configuration, the thrust, efficiency, and specific impulse are measured. 1-m downstream centerline point measurements of most probable ion energy and plasma potential are achieved with a retarding potential analyzer and emissive probe, respectively Analysis of data point to three separate electron termination paths that govern the thruster-tochamber coupling.

Facility Effects on Helicon Plasma Source Operation

argon. The plasma potential and ion energy distribution are measured in the argon plasma plume using an RF compensated emissive probe and a retarding potential analyzer (RPA) respectively. Plasma properties measured for these operating conditions are compared to the values and trends published by the University of Wisconsin-Madison (UW-Madison) where the MadHeX is characterized during operation in a distinctly different facility. Operation at 25 sccm showed an axial plasma potential drop of 71 V over an axial distance of 30 cm. Both the drop in potential and axial distance over which it occurred exceeded the values observed by UW-Madison. Plasma potential at a volumetric flow rate of 30 sccm increased a total of 25 V over a 400 W increase in RF forward power and 55 V over a decrease in source magnetic field strength of 260 G in VTF-2. At 25 sccm, the plasma potential increased 61 V at the thruster exit over the same magnetic field strength decrease, similar to trends observed by UW-Madison. Ion energy was noted to increase by 55 V over a 400 W increase in forward RF power for a volumetric flow rate of 30 sccm while changes of less than 20 V between operating points were observed for the 25 sccm case.