Design and Performance Analysis of a Low Current Hollow Cathode with LaB6 for Electric Propulsion on Satellite

Abstract

This research presents the design and experimental validation of a low-power lanthanum hexaboride (LaB₆) hollow cathode for satellite electric propulsion. Hollow cathodes are critical components in ion and Hall-effect thrusters, performing two essential functions: emitting electrons to ionize the propellant gas for plasma generation, and injecting electrons into the exhaust beam to neutralize the positively charged ion plume. Without neutralization, spacecraft charging would destabilize thrust and potentially damage onboard electronics. Conventional hollow cathodes utilize impregnated emitters porous metal structures infused with low-work function chemicals. Although effective, these emitters are highly sensitive to contamination from oxygen or moisture in the propellant, require controlled heating cycles with high power consumption, and exhibit limited durability under frequent on-off cycling required for small satellite station-keeping. LaB₆ is a refractory ceramic with high melting point and superior resistance to contamination, was investigated as an alternative emitter material. The cathode structure, designed using SolidWorks, incorporated a LaB₆ crystal within an emitter cavity approximately 30 mm in length. Critical geometric parameters included orifice diameter, which controls internal pressure and plasma flow, and keeper electrode optimization for electron extraction and internal component shielding. Electro-thermal simulations were conducted at multiple temperatures to characterize emission behavior. At 1300°C, current density distribution across the emitter surface was non uniform with edge current crowding, indicating localized overheating and uneven erosion that would reduce operational lifetime. At 1500°C, strong thermionic emission produced significant space-charge effects, where a dense electron cloud formed near the emitter surface, limiting further emission by opposing the applied electric field. This space-charge-limited regime demonstrated a self-regulating mechanism preventing uncontrolled current increase and thermal runaway. Electron trajectory analysis revealed complex motion influenced by electric and magnetic fields, leading to non-linear paths and localized heat deposition on nearby structures. At 1700°C, the system exhibited stable response reaching steady emission in under one millisecond, with current density sufficient for thruster operation while remaining below levels causing rapid material evaporation. Experimental validation was performed in a vacuum chamber through four design iterations. Initial graphite heater wire exhibited resistance instability at high temperature causing premature burnout. Second iteration employing niobium heater achieved successful plasma ignition but experienced insulation failure due to high-temperature degradation of ceramic insulating properties, particularly at graphite contact interfaces creating unintended conductive paths. Subsequent modifications introduced physical spacing between conductive elements and improved insulation coverage. Final configuration achieved stable plasma plume generation with currents up to 5 A, consistent with simulation predictions. Repeated ignition cycles were successful, demonstrating operational robustness suitable for small satellite applications. Minor ceramic surface degradation indicated areas for material improvement without compromising functionality. The results confirm LaB₆ as a viable emitter material for low-power hollow cathodes, with simulations accurately predicting emission behavior and thermal distribution. Key findings include the necessity of operating temperatures above 1500°C for uniform emission, the self regulating nature of space-charge-limited operation, and the critical importance of high temperature insulation materials and sub-millimeter assembly precision. Future work will focus on improved insulation materials and extended lifetime testing for long-duration space missions