Richard E. Wirz

Effects of Internally Mounted Cathodes on Hall Thruster Plume Properties

The effects of cathode position on the operation and plume properties of an 8-kW Hall thruster are discussed. Thruster operation was investigated at operating conditions ranging from 200 to 500 V of discharge voltage, 10-40 A of discharge current, and 2-8 kW of discharge power, with a cathode positioned either in the traditional externally mounted configuration outside the outer magnetic pole piece or in an internally mounted configuration central to the inner magnetic core. With the external cathode, substantial emission in the visible spectrum that follows magnetic field lines surrounds the exterior pole pieces of the thruster. With the internal cathode, the emission is largely absent while the cathode plume is compressed and elongated in the axial direction by the strong axial magnetic field on the thruster centerline. Discharge current oscillation and ion species fraction measurements were found to be similar for the cathode locations, whereas the operation with the internal cathode was found to favor an improved coupling of the cathode plume with the thruster discharge. Ion current density measurements show that with respect to externally mounted designs, internally mounted cathodes reduce plume divergence and increase the symmetry of the near-field plume. The impacts of internally mounted cathodes on thruster physics and spacecraft integration activities are assessed.

Plasma processes of DC ion thruster discharge chambers

This study reveals important aspects of discharge plasma processes and advances state-of-the-art discharge modeling. A multi-component hybrid 2-D computational Discharge Chamber Model (DCM) was developed to help identify important ion thruster discharge processes and aid in thruster design and optimization. The model accounts for the five major chamber design parameters (chamber geometry, magnetic field, discharge cathode, propellant feed, and ion extraction grid characteristics) and self-consistently tracks the effects of the four discharge plasma species (neutral propellant atoms, secondary electrons, primary electrons, and ions). Good agreement with NSTAR thruster data was found by weighting the influence of Bohm-type non-classical diffusion per the relative level of ion-centered electron collisions. According to DCM results, ion thrusters operate in an intermediately ionized plasma regime that is between fully and weakly ionized approximations. The model analyses show that the peak observed in the NSTAR beam profile is due to double ions that are created by over-confinement of primary electrons on the thruster axis. A design analysis performed with DCM shows that simply increasing the strength of the middle magnet ring of the NSTAR thruster leads to remarkable improvements in thruster efficiency and beam flatness. These improvements are due to the propensity of this modified magnetic field to evenly distribute the primary electrons, which decreases double ion content on-axis, increases primary electron utilization, and promotes beam flatness. Simply increasing primary confinement does not guarantee better performance. The modified design also closes a higher magnetic field contour (~40 Gauss, instead of ~30 Gauss) and thus shows better plasma confinement between the cusps.

Miniature Ion Thrusters for Precision Formation Flying

Miniaturized ion thrusters are attractive for precision formation flying and many other future space missions. Their potential for high-efficiency and low-noise operation provides the mission designer with a uniquely attractive option in the mN thrust range. Caltech and JPL have designed and tested a small, 3 cm diameter ion thruster that demonstrates propellant efficiency > 80% and Isp > 3100s using a cathode generated xenon discharge. These investigations have shown that desirable discharge confinement efficiency is possible with reasonably sized magnets, and consequently low overall thruster weight. Robust, long-life ion extraction grid geometries were designed using computational tools and have performed well during experimental testing. Several flight-worthy cathode designs are currently under investigation for the neutralizer and discharge cathodes. In particular, recent design efforts have focused on miniature hollow and direct emission cathodes. Direct emission cathodes using hexaboride emitters are being considered for their applicability to precision formation flying missions, fast cycling times, scalability to even smaller thruster sizes (<3cm), and ability to run with low-purity propellants. Early testing of a direct emission LaB6 cathode has shown that such a cathode can provide good overall performance if consistent cathode heater performance can be maintained. Future testing will focus on the design and validation of efficient, flight- worthy cathodes and demonstration of thruster lifetime.

Development and Initial Testing of a Magnetically Shielded Miniature Hall Thruster

The scaling of magnetically shielded Hall thrusters to low power is investigated through the development and fabrication of a 4-cm Hall thruster. During initial testing, the magnetically shielded miniature Hall thruster was operated at 275 V discharge voltage and 325-W discharge power. Inspection of the channel walls after testing suggests that the outer discharge channel wall was successfully shielded from high-energy ion erosion while the inner channel wall showed evidence of weaker shielding, likely due to magnetic circuit saturation. Scanning planar probe measurements taken at two locations downstream of the thruster face provided ion current density profiles. The ion current calculated by integrating these data was 1.04 A with a plume divergence half-angle of 30°. Swept retarding potential analyzer measurements taken 80-cm axially downstream of the thruster measured the most probable ion voltage to be 252 V. The total thruster efficiency was calculated from probe measurements to be 43% (anode efficiency of 59%) corresponding to a thrust of 19 mN at a specific impulse of 1870 s. Discharge channel erosion rates were found to be approximately three orders of magnitude less than unshielded Hall thrusters, suggesting the potential for a significant increase in operational life.

Experimental and computational investigation of the performance of a micro-ion thruster

A micro-ion thruster assembly with a characteristic diameter of 3-cm has been developed at JPL for testing and optimization of various system parameters.

Effects of magnetic field topography on ion thruster discharge performance

Traditional magnetic field design techniques for dc ion thrusters typically focus on closing a sufficiently high maximum closed magnetic contour, Bcc, inside the discharge chamber. In this study, detailed computational analysis of several modified NSTAR thruster 3-ring and 4-ring magnetic field geometries reveals that the magnetic field line shape as well as Bcc determines important aspects of dc ion thruster performance (i.e. propellant efficiency, beam flatness and double ion content). The DC-ION ion thruster model results show that the baseline NSTAR configuration traps the primary electrons on-axis, which leads to the high on-axis plasma density peak and high double ion content observed in experimental measurements. These problems are further exacerbated by simply increasing Bcc and not changing the field line shape. Changing the field line shape to prevent on-axis confinement (while maintaining the NSTAR baseline Bcc) improves thruster performance, improves plasma uniformity and lowers double ion content. For these favorable field line geometries, we observe further improvements to performance with increased Bcc, while maintaining plasma uniformity and low double ion content. These improvements derive from the fact that the field lines guide the high-energy primaries to regions where they are most efficiently used to create ions while a higher Bcc prevents the loss of ions to the anode walls. Therefore, it is recommended that the ion thruster designer first establish a divergent field line shape that ensures favorable beam flatness, low double ion content and reasonable performance; then the designer may adjust the Bcc to attain desirable performance and stability for the target discharge plasma conditions.

Hollow Cathode and Low-Thrust Extraction Grid Analysis for a Miniature Ion Thruster

Miniature ion thrusters are well suited for future space missions that require high efficiency, precision thrust, and low contamination in the mN to sub-mN range. JPLs miniature xenon Ion (MiXI) thruster has demonstrated an efficient discharge and ion extraction grid assembly using filament cathodes and the internal conduction (IC) cathode. JPL is currently preparing to incorporate a miniature hollow cathode for the MiXI discharge. Computational analyses anticipate that an axially upstream hollow cathode location provides the most favorable performance and beam profile; however, the hot surfaces of the hollow cathode must be sufficiently downstream to avoid demagnetization of the cathode magnet at the back of the chamber, which can significantly reduce discharge performance. MiXIs ion extraction grids are designed to provide > 3 mN of thrust; however, previous to this effort, the low-thrust characteristics had not been investigated. Experimental results obtained with the MiXI-II thruster (a near replica or the original MiXI thruster) show that sparse average discharge plasma densities of ∼ 5 × 10 15 – 5 × 10 16 m - 3 allow the use of very low beamlet focusing extraction voltages of only ∼ 250 –500 V, thus providing thrust levels as low as 0.03 mN for focused beamlet conditions. Consequently, the thrust range thus far demonstrated by MiXI in this and other tests is 0.03–1.54 mN.

CubeSat Lunar Mission Using a Miniature Ion Thruster

The feasibility of CubeSats utilizing the Miniature Xenon Ion (MiXI) thruster for lunar missions is the focus of this investigation. The successful heritage, simplicity, and low cost of CubeSats make them attractive candidates for scientific missions to the Moon. This investigation presents the first-order design process for developing a lunar mission CubeSat. The results from this process are then applied to a 3U CubeSat equipped with a MiXI thruster and specifically designed to reach the lunar surface from a low Earth orbit. The small, 3cm diameter MiXI thruster utilized is capable of producing 0.1–1.553mN of thrust with a specific impulse of over 3000s and is projected to be capable of generating over

Qualification of Commercial XIPS ® Ion Thrusters for NASA Deep Space Missions

‡‡ Electric propulsion systems based on commercial ion and Hall thrusters have the potential for significantly reducing the cost and schedule-risk of Ion Propulsion Systems (IPS) for deep space missions. The large fleet of geosynchronous communication satellites that use SEP, which will approach 40 satellites by year-end, demonstrates the significant level of technical maturity and spaceflight heritage achieved by the commercial electric propulsion systems. A program to delta-qualify XIPS ® ion thrusters for deep space missions is underway at JPL. This program includes modeling of the thruster grid and cathode life, environmental testing of a 25-cm EM thruster over anticipated vibe and thermal requirements for deep space missions, and wear testing of the thruster cathodes to demonstrate the life and benchmark the model results. This paper will present the deltaqualification status of the XIPS thruster and discuss the life and reliability with respect to known failure mechanisms.

Ion–Neutral Collision Modeling Using Classical Scattering With Spin-Orbit Free Interaction Potential

A particle-in-cell Monte Carlo collision model is developed to explore dominant collisional effects on high-velocity xenon ions incident to a quiescent xenon gas at low neutral pressures. The range of neutral pressure and collisionality examined are applicable for electric propulsion as well as plasma processing devices; therefore, the computational technique described herein can be applied to more complex simulations of those devices. Momentum and resonant charge-exchange collisions between ions and background neutrals are implemented using two different models, classical scattering with spin-orbit free potential and variable-hard-sphere model, depending on the incident particle energy. The primary and charge-exchange ions are tracked separately, and their trajectories within a well-defined “Test Cell” domain are determined. Predicted electrode currents as a function of the Test Cell pressure are compared with electrode currents measured in an ion gun experiment. The simulation results agree well with the experiment up to a Test Cell pressure corresponding to a mean free path of the Test Cell length and then start to deviate with increasing collisionality at higher pressures. This discrepancy at higher pressures is likely due to the increasing influence of secondary electrons emitted from electrodes due to the high-velocity impacts of heavy species (i.e., beam ions and fast neutrals created by charge-exchange interaction) at the electrode surfaces.