Abstract
Lanthanum hexaboride hollow cathodes represent a viable option for high power Hall effect thruster applications, under development for the next generation of manned and robotic interplanetary missions. In this scenario, SITAEL and the University of Pisa are actively developing high current hollow cathodes capable of providing discharge current in the range 10-100 A to be coupled with high power Hall effect thrusters. The cathode design is based on an in-house theoretical model of the internal sections of the cathode, recently integrated with a simplified model of the cathode plume. Despite the application of hollow cathodes on flight and laboratory model Hall effect thrusters, many questions remain unsolved. In particular, issues related to onset of instabilities, due to plume mode or ion acoustic turbulence, are still unclear, while it is known that they can affect the overall performance of the cathode and thruster unit. This paper focuses on the experimental investigation of the cathode plume by means of measurements of the main plasma parameters, at different operating conditions and for different cathode geometry. Two cathodes were investigated, namely HC20 and HC60, designed to be coupled with SITAEL’s HT5k and HT20k (5 kW- and 20kW-class) Hall effect thrusters. The cathodes were mounted in stand-alone configuration with an auxiliary cylindrical anode. The experimental campaign was performed using triple Langmuir probes as plasma diagnostic system. The probes were mounted on scanning mechanisms to measure the plume parameters at various radial and axial distances from the keeper exit. General trends of electron temperature, plasma potential and plasma density are reported in terms of discharge current, mass flow rate and cathode orifice geometry. The results highlight that the cathode plate orifice selection affects the plume mode onset, giving the possibility to extend the stable mode of cathode operation in the current range required by the thruster.
Highlights
High power space electric propulsion has increased its attractive perspective for the generation exploration missions and for transportation of large space tugs toward future space stations [1]
The experimental apparatus includes a cylindrical anode made of stainless steel and the cathode under test, both placed over a base plate, insulated with appropriate ceramic supports
The plasma parameters for HC20 (Figure 6, right column) show that plume mode of operation is characterized by a 30% higher electron temperature and 40% lower plasma potential
Summary
High power space electric propulsion has increased its attractive perspective for the generation exploration missions and for transportation of large space tugs toward future space stations [1]. At the end of both tests the keeper electrode was completely eroded, exposing the cathode orifice plate to the external discharge plasma These results have been attributed to high-energy ions bombarding and sputtering the external surfaces of the cathode. HC20 can reliably provide 5–30 A at a mass flow rate lower than 2.5 mg/s (Xe) Both cathodes were previously tested in stand-alone configuration to characterize their performance as well as in coupled mode with the thrusters [32]. The heater is turned on for 10 min at 300 W of power so that the emitter temperature is increased to allow for the thermionic emission In this way, after the initialization of the mass flow rate, the ignition can be performed by applying only the discharge voltage between the cathode and the keeper required to ionize the propellant.
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