James E. Polk

Hollow cathode theory and experiment. II. A two-dimensional theoretical model of the emitter region

Despite their long history and wide range of applicability that includes electric propulsion, detailed understanding of the driving physics inside orificed hollow cathodes remains elusive. The theoretical complexity associated with the multicomponent fluid inside the cathode, and the difficulty of accessing empirically this region, have limited our ability to design cathodes that perform better and last longer. A two-dimensional axisymmetric theoretical model of the multispecies fluid inside an orificed hollow cathode is presented. The level of detail attained by the model is allowed by its extended system of governing equations not solved for in the past within the hollow cathode. Such detail is motivated in part by the need to quantify the effect(s) of the plasma on the emitter life, and by the need to build the foundation for future modeling that will assess erosion of the keeper plate. Results from numerical simulations of a 1.2-cm-diam cathode operating at a discharge current of 25A and a gas flow ra...

Numerical simulations of ion thruster accelerator grid erosion

The highly successful demonstration of ion propulsion on Deep Space 1 has stimulated the study of more demanding applications of this technology. These future applications require ion thrusters capable of providing significantly greater specific impulses and total impulses than the current state-of-the-art. Higher specific impulses aggravate the known wear out mechanisms of the ion accelerator system. Computer simulations of the ion accelerator system operation and erosion are essential tools for the development of ion thrusters to meet the demand for higher specific impulse and longer life. Two-dimensional and three-dimensional computer codes have been developed at JPL and are used to provide insight into the processes limiting the life of the accelerator grid. The 2D code described herein was used to identify a key feature of operation at high Isp. That is, as the beam voltage is increased the energy of the charge-exchange ions hitting the hole walls increases in proportion to the beam voltage and more rapidly than the increase in the magnitude of the accelerator grid voltage. This has serious implications for the design of long-life, high specific impulse thrusters.

Validation of the NSTAR Ion Propulsion System on the Deep Space One Mission: Overview and Initial Results

Deep Space 1 is the first interplanetary spacecraf t to use an ion propulsion system for the p r imary delta-v maneuvers. The purpose of the mission is to validate a number of technologies, including ion propulsion and a high degree of spacecraft autonomy, on a fiyby of an asteroid and two comets. The ion propulsion system has operated now for a total of 1791 hours at engine power levels ranging from 0.48 to 1.94 kW and has completed the deterministic thrusting requi red for an encoun te r w i th the asteroid 1992KD in late July, 1999. The system has worked extremely well after an initial grid short was cleared after launch. O p eration during this primary mission phase has demonstrated all ion propuls ion system and au tonomous navigation functions. All propulsion system operating parameters are ve ry close to the expected values wi th the excep tion of the thrust at higher power levels, which is about 2 percent lower than calculated values. This paper provides an overview of the system and presents the first flight validation data on an ion propulsion system i n interplanetary space.

Wear Mechanisms in Electron Sources for Ion Propulsion, 2: Discharge Hollow Cathode

The wear of the keeper electrode in discharge hollow cathodes is a major impediment to the implementation of ion propulsion onboard long-duration space science missions. The development of a predictive theoretical model for hollow cathode keeper life has long been sought, but its realization has been hindered by the complexities associated with the physics of the partially ionized gas and the associated erosion mechanisms in these devices. Thus, although several wear mechanisms have been hypothesized, a quantitative explanation of life test erosion profiles has remained incomplete. A two-dimensional model of the partially ionized gas in a discharge cathode has been developed and applied to understand the mechanisms that drove the erosion of the keeper in two long-duration life tests of a 30-cm ion thruster. An extensive set of comparisons between predictions by the numerical simulations and measurements of the plasma properties and of the erosion patterns is presented. It is found that the near-plume plasma oscillations, predicted by theory and observed by experiment, effectively enhance the resistivity of the plasma as well as the energy of ions striking the keeper.

Three-dimensional particle simulations of ion-optics plasma flow and grid erosion

A fully three-dimensional computer particle simulation model for ion optics is developed. This model allows multiple apertures to be included explicitly in the simulation domain and determines the upstream sheath and downstream beam neutralization through simulations. Simulations are performed for the NSTAR ion-thruster optics, and results are compared with grid-erosion measurements obtained during NSTAR long-duration test. It is shown that the simulation not only predicts accurately all of the features in the measured erosion pattern but also gives excellent quantitative agreement with the measured erosion depth.

An Overview of the Nuclear Electric Xenon Ion System (NEXIS) Activity

The Nuclear Electric Xenon Ion System (NEXIS) research and development activity within NASAs Project Prometheus, was one of three proposals selected by NASA to develop thruster technologies for long life, high power, high specific impulse nuclear electric propulsion systems that would enable more robust and ambitious science exploration missions to the outer solar system. NEXIS technology represents a dramatic improvement in the state-of-the-art for ion propulsion and is designed to achieve propellant throughput capabilities >= 2000 kg and efficiencies >= 78% while increasing the thruster power to >= 20 kW and specific impulse to >= 6000 s. The NEXIS technology uses erosion resistant carbon-carbon grids, a graphite keeper, a new reservoir hollow cathode, a 65-cm diameter chamber masked to produce a 57-cm diameter ion beam, and a shared neutralizer architecture to achieve these goals. The accomplishments of the NEXIS activity so far include performance testing of a laboratory model thruster, successful completion of a proof of concept reservoir cathode 2000 hour wear test, structural and thermal analysis of a completed development model thruster design, fabrication of most of the development model piece parts, and the nearly complete vacuum facility modifications to allow long duration wear testing of high power ion thrusters.

Characterization of cathode keeper wear by surface layer activation

In this study, the erosion rates of the discharge cathode keeper in a 30 cm NSTAR configuration ion thruster were measured using a technique known as Surface Layer Activation (SLA). This diagnostic technique involves producing a radioactive tracer in a given surface by bombardment with high energy ions. The decrease in activity of the tracer material may be monitored as the surface is subjected to wear processes and correlated to a depth calibration curve, yielding the eroded depth. Analysis of the activities was achieved through a gamma spectroscopy system. The primary objectives of this investigation were to reproduce erosion data observed in previous wear studies in order to validate the technique, and to determine the effect of different engine operating parameters on erosion rate. The erosion profile at the TH 15 (23 kw) setting observed during the 8200 hour Life Demonstration Test (LDT) was reproduced. The maximum keeper erosion rate at this setting was determined to be 0.085 pm/hr. Testing at the TH 8 (1.4 kw) setting demonstrated lower erosion rates than TH 15, along with a different wear profile. Varying the keeper voltage was shown to have a significant effect on the erosion, with a positive bias with respect to cathode potential decreasing the erosion rate significantly. Accurate measurements were achieved after operating times of only 40 to 70 hours, a significant improvement over other erosion diagnostic methods.

Asteroid Return Mission Feasibility Study

This paper describes an investigation into the technological feasibility of finding, characterizing, robotically capturing, and returning an entire Near-Earth Asteroid (NEA) to the International Space Station (ISS) for scientific investigation, evaluation of its resource potential, determination of its internal structure and other aspects important for planetary defense activities, and to serve as a testbed for human operations in the vicinity of an asteroid. Reasonable projections suggest that several dozen candidates NEAs in the size range of interest (~2-m diameter) will be known before the end of the decade from which a suitable target could be selected. The conceptual mission objective is to return a ~10,000-kg asteroid to the ISS in a total flight time of approximately 5 years using a single Evolved Expendable Launch Vehicle. Preliminary calculations indicate that this could be accomplished using a solar electric propulsion (SEP) system with high-power Hall thrusters and a maximum power into the propulsion system of approximately 40 kW. The SEP system would be used to provide all of the post-launch ∆V. The selected asteroid would have an unrestricted Earth return Planetary Protection categorization, and would be curated at the ISS where numerous scientific and resource utilization experiments would be conducted. Asteroid material brought to the ground would be curated at the NASA Johnson Space Center. This preliminary study identified several areas where additional work is required, but no show stoppers were identified for the approach that would return an entire 10,000-kg asteroid to the ISS in a mission that could be launched by the end of this decade.

NSTAR flight thruster qualification testing

Deep Space 1 is a technology demonstration mission scheduled to be launched in October 1998. One of those technologies is the NSTAR 30 cm diameter xenon ion thruster which will provide the primary propulsion. Three Flight-design thrusters were designed and built by Hughes Electron Dynamics Division, with assistance from NASAs Lewis Research Center. The first thruster was a Pathfinder to finalize the fabrication and assembly procedures for the other thrusters. Two flight-worthy thrusters were then fabricated and tested to Protoflight Qualification levels at NASAs Lewis Research Center and Jet Propulsion Laboratory. Each thruster was performance tested before and after Vibration Tests, integrated with different flight power processors and digital control interface units, and underwent Thermal Vacuum Tests with engine starts from -97 °C. Performance tests included neutralizer, discharge chamber, and ion optics characterizations as well as measurements of thruster efficiency over the full 0.5 to 2.3 kW power throttle range. The performance, at both component and thruster levels, was as expected and found to be quite repeatable with negligible dispersion between thrusters. After final functional tests, one thruster was installed on the DS 1 spacecraft while the other was set aside as a flight spare.

Cathode phenomena in plasma thrusters

For the optimization of cathodes of steady-state plasma accelerators, an extensive experimental program has been carried out at IRS to investigate the erosion mechanisms. On different thruster types and with an additional fundamental experiment, the influences of geometry, chamber pressure, amount of ThO2, cathode temperature, power, and gas species were explored. Propellants Ar, H2, N2 were used which had been purified with a special filter to remove residual water and oxygen. The erosion rates during the ignition phase as well as during continuous operation were measured. The temporal evolution of the cathode erosion rate was measured by surface layer activation. At high current densities within the cathode, severe problems were encountered with all thruster geometries. The cathodes developed cracks and started to melt. To clarify the reason for this behavior, some cathodes were investigated metallurgically.