Vadim Khayms

Demonstration of 10,400 Hours of Operation on a 4.5 kW Qualification Model Hall Thruster

Between 2007 and 2009, Aerojet and Lockheed Martin Space Systems Company (LMSSC) successfully extended the demonstrated operating duration of the qualification model BPT-4000 4.5 kW Hall thruster beyond 10,400 hrs. A total of 452 kg of xenon were consumed during the entire qualification marking the most throughput ever demonstrated on a Hall thruster. The BPT-4000 Hall thruster is part of a 4.5 kW Hall Thruster Propulsion System (HTPS) developed jointly by Aerojet and LMSSC. The system is slated for initial launch in the summer of 2010 on the first Advanced EHF spacecraft. The testing demonstrated the thruster’s ability to provide more than 8.7 MN-s of total impulse and 7,316 ignition cycles. At the conclusion of testing, the thruster showed no signs of degradation in performance and all health indicators were stable. Most significantly, there was no measurable insulator ring erosion from 5,600 hrs to 10,400 hrs indicating that the thruster had reached a “zero” erosion configuration. This result demonstrates that Hall thrusters can, if designed properly, achieve lifetimes comparable to ion thrusters. It also eliminates one of the perceived barriers to the use of Hall thrusters for applications such as asteroid fly-by, cargo transfer, and satellite servicing missions that require significant throughput.

Direct Measurement of Axial Momentum Imparted by an Electrothermal Radiofrequency Plasma Micro-Thruster

Gas flow heating using radio frequency plasmas offers the possibility of depositing power in the centre of the flow rather than on the outside, as is the case with electro-thermal systems where thermal wall losses lower efficiency. Improved systems for space propulsion are one possible application and we have tested a prototype micro-thruster on a thrust balance in vacuum. For these initial tests, a fixed component radio frequency matching network weighing 90 grams was closely attached to the thruster in vacuum with the frequency agile radio frequency generator power being delivered via a 50 Ohm cable. Without accounting for system losses (estimated at around 50~

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

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Status of Hall Thruster Integration Activities at Lockheed Martin Space Systems Company

), for a few 10s of Watts from the radio frequency generator the specific impulse was tripled to

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

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Long Term Exposure of a Biased Solar Array to a Hall Thruster Plume

48 seconds and the thrust tripled from 0.8 to 2.4 milli-Newtons.

4.5 KW Hall Thruster System Qualification Status

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.

Ion flux, energy, and charge-state measurements for the BPT-4000 Hall thruster

In an effort to increase payload capability of future commercial and military spacecraft, a program was initiated at Lockheed Martin several years ago to investigate Hall Current Thruster (HCT) technology. Lockheed Martin and Aerojet (formerly General Dynamics) entered into a long-term agreement in the Fall of 2000 to develop and qualify flight versions of the BPT-4000 Hall thruster, a power processing unit (PPU), and a xenon flow controller (XFC). In parallel with the development of flight hardware, Lockheed Martin has been pursuing a number of internal integration studies. The main emphasis of this comprehensive Hall Thruster Propulsion System (HTPS) development and integration program has been to gain quantitative information on the effects of the thruster plume on various spacecraft subsystems and mitigate risk associated with the introduction of this new propulsion technology on future spacecraft. Integration of Hall thrusters requires a detailed understanding of the potential effects induced by the thruster on satellite operations. Issues related to the thruster plume include: sputtering of spacecraft surfaces due to ion impingement, re-deposition of sputtered materials onto critical surfaces and surface contamination, transmission of bus and payload signals through the thruster plume, electromagnetic interference due to thruster radiated and conducted emissions, surface charging and electrostatic discharge (ESD) susceptibility, the effect of erosion/contamination on solar array performance, and a number of other concerns. This paper presents some of the major highlights of this effort and summarizes the work completed to date.

Electrical Facility Effects on Hall Thruster Cathode Coupling: Performance and Plume Properties

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.

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

This paper summarizes nhe results of an experiment in which flight-like solar array coupons and a mockup coupon were exposed to a Hall thruster plume in a biased configuration representative of a typical arrangement achieved by the strings of solar cells in a CEO environment during thruster firing. The results of this test helped validate the model used to make erosion predictions at the spacecraft level, to understand the effects of interconnect bias on material erosion and localized re-deposition, and to evaluate the overall impact of the combined plasma effects on the end-of-life performance of solar arrays in orbit.