Reduced-order particle-in-cell simulations of a high-power magnetically shielded Hall thruster

High-power magnetically shielded Hall thrusters have recently emerged to meet the needs of the next-generation space missions. Even though a few such thrusters are currently undergoing their late-stage development campaigns, many unanswered questions yet exist concerning the behavior and evolution of the plasma in these large-size thrusters that feature an unconventional magnetic field topology. Noting the complex, multi-dimensional nature of plasma processes in Hall thrusters, high-fidelity particle-in-cell simulations are optimal tools to study the intricate plasma behavior. Nonetheless, the significant computational cost of traditional PIC schemes renders simulating high-power thrusters without any physics-altering speed-up factors unfeasible. Thus, in this article, we demonstrate the applicability of the novel reduced-order PIC scheme for a cost-efficient, self-consistent study of the high-power Hall thrusters by performing simulations of a 20 kW magnetically shielded Hall thruster along the axial-azimuthal and radial-azimuthal coordinates. The axial-azimuthal simulations are performed for three operating conditions in a rather simplified representation of the thruster's inherently 3D configuration. Nevertheless, we resolved self-consistently an unprecedented 650 us of the discharge evolution without any ad-hoc electron mobility model, capturing several breathing cycles and approximating the experimental performance parameters with an accuracy of 70 to 80 % across the operating conditions. The radial-azimuthal simulations casted further light on the evolution of the azimuthal instabilities and the resulting variations in the electrons' cross-field mobility and the plasma-wall interactions. Particularly, we observed the development of a long-wavelength, relatively low-frequency wave mode near the exit plane of the thruster's channel that induces a notable electron transport.Comment: 29 pages, 25 figure

Mars- plus Europa-INPPS Flagship Missions with High Power Electric Thrusters and Heavy Science Payload

2020/2021 orbit calculations (DLR Bremen + MAI Moscow): 3 Orbits for one non-human MARS/EUROPA-INPPS flagship 1. Orbit Earth (11 Oct 2026) => Mars (28 Jan 2028); 2. Orbit Mars (22 Sept 2028) => Earth (29 Jul 2029); 3. Orbit Earth (12 Oct 2031) => Jupiter / Europa (6 Dec 2035) 22 ETs (for the INPPS flagship CET): example - gridded CET plate with combined ET

Non-Human + Human Deep Space Exploration: By The NEP INPPS Flagship

INPPS Flagship Structure & Subsystems, Ground Based Tests, Pure NEP, International Electric Thrusters, Humans Internationally To Mar

Humans to Mars: by MARS- plus EUROPA-INPPS Flagship Mission

The first non-human INPPS (International Nuclear Power and Propulsion System) flagship flight with orbits Earth-Mars-Earth-Jupiter/Europa (after 2025) is the most maximal space qualification test of INPPS flagship to carry out the second INPPS flagship flight to Mars with humans (in the 2030th). This high power space transportation tug is realistic because of A) the successful finalization of the European-Russian DEMOCRITOS and MEGAHIT projects with their three concepts of space, ground and nuclear demonstrators for INPPS realization (reached in 2017), B) the successful ground based test of the Russian nuclear reactor with 1MWel plus important heat dissipation solution via droplet radiators (confirmed in 2018), C) the space qualification of the Russian reactor by 2025 and D) the perfect celestial constellation for a Earth-Mars/Phobos-Earth-Jupiter/Europa trajectory between 2026 and 2035. Therefore the talk sketches the preparation status of INPPS flagship with its subsystems. Critical performance will be studied by parallel realizations of the ground and nuclear demonstrators of DEMOCRITOS (until 2025). The space qualification of INPPS with all subsystems including the nuclear reactor in the middle of the 2020th plus the INPPS tests for about one to two years - first in high Earth orbit robotic assembly phase of INPPS and later extended in nearby Earth space environment flight - means a complete concepts driven approval for all applied INPPS space subsystem technologies. It is also important to consider wider aspects for the overall mission implementation phase. Component like the nuclear reactor as the power source for the propulsion system will have to agree with the 1992 UN principles relevant to the use of nuclear power sources (NPS) in outer space. Therefore this talk will look into the legal and policy issues of nuclear space systems related to the international realization of mission design, requirements of associated safety regulations (including AI applications in the subsystems) and new aspects for INPPS flagship commercialization and new media communication on board

Modeling and Scaling of Plasma Thrusters

The name "Plasma Thrusters" usually adresses all those devices which generate thrust for space propulsion using a working fluid that remains in the state of plasma throughout the acceleration process. The present work started with the aim of carrying out a broad investigation of plasma flows inside the acceleration channels of plasma thrusters, paying special attention to the two best known devices belonging to this group: Hall Effect Thrusters (HET) and Magnetoplasmadynamic Thrusters (MPDT). Eventually, the work ended up focusing on two main issues, one related to Hall thrusters and the other one related to MPD thrusters. The thesis has been thus split into two main parts, the first one concerning the problem of defining a proper scaling methodology for Hall Effect thrusters and the second one devoted to numerical investigations of a plasma flow inside an MPD thruster. About the first topic, a robust scaling methodology for HETs has been developed starting from the physical laws which govern the plasma behaviour inside the acceleration chamber. A reliable scaling method is a valuable tool in the preliminary design of HETs, because it allows the designer to immediately have a taste of what the thruster geometric and operational parameters will be for a chosen power level (or for achieving a desired performance). The method object of this thesis has a great flexibility and can be easily adapted to work satisfactorily both at high and low power levels. The extension of its applicability to low power levels has been possible exploiting Alta's wide experience on a low power HET like the HT-100 and with the aid of a dedicated numerical code describing the plasma flow along the chamber. The scaling method here proposed has been tested versus a large number of experimental data and has proved to work well in a wide range of conditions, not only for different power levels but also for different thruster configuration(the method has been tested both on conventional Hall thrusters and on thrusters in "anode layer" configuration). Unfortunately, one of the most important parameters to be considered when designing a new thruster, the operational lifetime, cannot be accurately predicted using the relatively simple equations involved in the scaling method. The channel erosion process which determines the lifetime of a Hall thruster is a complex phenomenon and has to be treated separately. For this purpose a considerable effort has been devoted to develop a numerical method able to predict the erosion rate during the thruster operation. The method is based on a Monte Carlo Collision (MCC) simulation of the ion impacts with the walls and it has been coupled with the scaling method (whose output, in terms of thruster geometric and operational parameters, constitutes the input for the numerical simulation of the channel erosion). The combination of the scaling method and the numerical code for HET lifetime prediction turns out to be an effective instrument for designing new Hall effect thrusters. The second main topic analyzed during this work on plasma thrusters concerns MPD thrusters and consisted in the development of a numerical method to investigate the behaviour of plasma flows inside these devices. MPD thrusters have always been plagued by plasma instability phenomena, which occur for high intensities of the applied current. These instabilities prevent smooth flow acceleration and severely affect the overall performance. Understanding the mechanism lying behind their occurrence is crucial to counter their detrimental effects and improve the thruster efficiency. Here, as a first step, a 1D model based on the equations of magnetohydrodynamics has been developed. The model, as well as the numerical code which comes with it, is very simple but in spite of its simplicity it successfully points out several interesting aspects. It describes the different plasma acceleration mechanisms in an MPD thruster, giving an estimate of the thruster performance and a rough prediction of the instability threshold (in terms of applied discharge current). However, such a simple model is not able to give any description of flow behaviour after the instability occurrence, nor it can be used for describing a flow in complex duct geometries. To have a more powerful tool for attempting a description of what happens in real plasma flows, a 3D code which solves the full magneto-fluid-dynamics (MFD) equations has been developed. MFD equations consist of the Navier-Stokes equations plus the Maxwell equations which describe the evolution of the electromagnetic fields. The 3D code presented in this thesis is based on an existing solver for what concerns the solution of the Navier-Stokes equations (which have been properly modified to take into account the electromagnetic contributions in the momentum and energy conservation equations). For what concerns the Maxwell equations, an independent code has been developed for solving them in space and time. Then, a strategy for coupling the two sub-codes has been defined and simple preliminary tests have been carried out to validate the full 3D code

Development of Hollow Cathodes for Space Electric Propulsion at Sitael

Hollow cathodes are electron sources used for the gas ionization and the beam neutralization in both ion and Hall effect thrusters (HETs). A reduction of power and propellant consumption from the cathode is particularly needed in small satellite applications, where power and mass budgets are inherently limited. Concurrently, the interest in high-power HETs is increasingly fostered for a number of space applications, including final positioning and station-keeping of Geostationary Earth Orbit (GEO) satellites, spacecraft transfers from Low Earth Orbit (LEO) to GEO, and deep-space exploration missions. As such, several hollow cathodes have been developed and tested at Sitael, each conceived for a specific power class of thrusters. A numerical model was used during the cathode design to define the geometry, in accordance with the thruster unit specifications in terms of discharge current, mass flow rate, and lifetime. Lanthanum hexaboride (LaB6) hollow cathodes were successfully developed for HETs with discharge power ranging from 100 W to 20 kW. Experimental campaigns were carried out in both stand-alone and coupled configurations, to verify the operation of the cathodes and validate the numerical model. The comparison between experimental and theoretical results are presented, offering a sound framework to drive the design of future hollow cathodes

SITAEL HC1 Low-Current Hollow Cathode

SITAEL is active in the field of electric propulsion and is involved in the development of different thruster technologies—mainly Hall thrusters (HTs)—of power levels ranging from 100 W up to 20 kW. Low-power HTs are the most effective choice to perform orbit transfer, drag compensation, and de-orbiting maneuvers for small satellites. This paper is dedicated to the activities regarding HC1, the hollow cathode conceived for the 100-W-class Hall thruster under development at SITAEL. Successful test campaigns were performed and are described, with emphasis on the improvements in the cathode design after an extensive research and development phase. The results are presented and discussed, along with future developments of the ongoing activities

Hall Effect Thruster Direct Drive PPUs, Experimental Investigation of the Cathode Potential Grounding Problem

In order to maximize system mass and cost savings, a Hall Effect Thruster Direct Drive Approach shall ideally get away with the need to provide galvanic isolation for the Anode voltage. In this respect, we present some experimental data aiming at characterising the implications of a direct connection between the HET Cathode Return Potential and the Spacecraft Ground