Electric Propulsion Space Research Articles

Satellite surface charging in LEO with ProPIC

ProPIC is a fully kinetic particle-in-cell (PIC) solver developed for space electric propulsion. This work has extended its capabilities to simulate satellite surface charging and wake generation in low Earth orbit (LEO). A novel scaling approach has been implemented, decreasing computational cost by more than one order of magnitude. The methodology and scaling approach have been verified against the revised orbital-motion-limited theory. The surface charging and wake generation in LEO have been examined for a satellite that is more complex and larger than what is typically handled with a fully kinetic PIC approach in LEO, particularly due to the presence of large solar panels. Notably, the simulated wake can be used to identify the optimal position for the plasma diagnostic sensor that minimizes interference with the wake. Moreover, despite not being a failure risk, the attitude greatly influences the surface charging of a satellite with large solar arrays installed parallel to the satellite speed vector. The study suggests that, for high positive pitch angles (>45∘), the surface charging of the solar panels can increase by as much as 75% compared to low negative pitching cases. Additionally, the study highlights that the pitch angle and satellite envelope along the motion direction significantly influence the potential gradients on the solar panels.

Multi-objective design optimization and physics-based sensitivity analysis of field emission electric propulsion for CubeSat platforms

Field-emission electric propulsion is an electrostatic space electric propulsion technology that offers various advantageous features including efficient design, high specific impulse, and versatile thrust capabilities ranging from micro-Newton to milli-Newton levels. These characteristics make this type of propulsion a promising technology for small satellite platforms, enabling precise attitude control, orbit maintenance, and de-orbiting through ionization and acceleration of a liquid metal propellant. The growing demand for small propulsion systems in CubeSat platforms has spurred significant progress in modeling and characterizing field emission electric propulsion thrusters to enhance their overall performance. However, little study has been conducted to investigate the effect of geometric configurations on electric fields or expelled ion trajectories for design optimization. In this study, multi-objective design optimization is performed by incorporating electrostatic simulation coupled with an analytical performance model into evolutionary algorithms based on prediction from surrogate modeling, aiming to optimize the thruster emission design to maximize thruster performance. Physical insights into the key design factors influencing the performance of field emission electric propulsion have been gained by probing into the interaction between ion particles and electric field behavior within the thruster. It has been found that the length of the emitter tip has a significant effect on plume divergence, i.e., a longer emitter tip under the influence of electric field at higher emitter current tends to result in lower initial acceleration of emitted ions and subsequently wider spread or divergence of the ion beam. A shorter emitter tip, on the other hand, generates a sharper E-field gradient, resulting in a more focused and narrower ion beam. Additionally, sensitivity analysis has identified the mass flow rate and potential distributions as the most influential design factors on performance due to the active roles they play in the performance generation process.

Unexpected Electrical Breakdown Characteristics and Protection of PPU in Space Electric Propulsion System

Unexpected Electrical Breakdown Characteristics and Protection of PPU in Space Electric Propulsion System

Ion–surface interactions in plasma-facing material design

A multi-scale simulation framework for ion–solid interactions in plasma-exposed materials provides crucial insight into advancing fusion energy and space electric propulsion. Leveraging binary-collision approximation (BCA) simulations, the framework uniquely predicts sputter yields and analyzes material transport within volumetrically complex materials. This approach, grounded in the validated BCA code TRI3DYN, addresses key limitations in existing models by accurately capturing ion–solid interaction physics. A case study is presented, highlighting the framework’s ability to replicate experimental sputter yield results, underscoring its reliability and potential for designing durable materials in harsh plasma environments. Insights into sputtering transport phenomenology mark a significant advancement in material optimization for improved resilience in plasma-facing applications.

Electric Propulsion in Space: Technology and Geopolitics

ABSTRACT During the 20th Century and within the first decade of the 21st, electric propulsion (EP) has played an important role in space activities, and its importance has not ceased to increase significantly in the last 15 years. The potential benefits of using EP on a bigger scale could be immense, reducing the cost of future space missions. The main hypothesis of this paper is that EP will play an increasingly important role in the system of international space relations in the coming decades with deep scientific, commercial, and geopolitical implications. To prove this, we examine the historical evolution, relevance, and future applications of such a propulsion system. Although there is extensive technical literature on EP, only a few works analyze electric space propulsion from an interdisciplinary perspective. On the contrary, much of the available information only provides an atomized, unipolar, and fragmented approach to the phenomenon. To fill this gap, this study presents an integrated analysis of the evolution, relevance, and trends of EP technologies, as well as their geopolitical implications in the space domain. The results obtained constitute a crucial contribution to the analysis of space propulsion from an interdisciplinary – socio-technological – perspective.

A Methodology for the Preliminary Design of a High-Efficiency Multistage Plasma Thruster

Space electric propulsion represents a class of power-limited systems that utilize the interaction of electromagnetic fields with ionized inert gas propellants to generate thrust. This technology has emerged as a highly fuel-efficient and sustainable alternative to chemical propulsion systems, particularly for satellite constellations. However, the miniaturization potential of EP systems is impeded by certain limitations, necessitating the exploration of novel architectures. The high-efficiency multistage plasma thruster (HEMP-T) stands as a promising contender for stand-alone missions due to its employment of a cusped magnetic-field topology, which effectively mitigates plasma-wall interactions and enhances overall efficiency even at low thrust levels. Despite the growing interest in HEMP-Ts, there is a dearth of comprehensive and streamlined preliminary design procedures for these systems. Prior research has predominantly focused on extensive numerical analyses, neglecting the development of efficient and accessible design tools. To bridge this gap, this paper presents a novel preliminary design tool derived from integrating established analytical models available in the literature. The proposed design tool also incorporates an iterative procedure that refines geometric properties using a 2D magnetostatic solver. Through the application of this tool, a 4 mN HEMP thruster was analyzed. This finally exhibited a specific impulse of approximately 2000s and a good efficiency level of 23%. Also, the results obtained for a 10 mN application align closely with those achieved by other types of EP thrusters.

Modelling of Cryopumps for Space Electric Propulsion Usage

Electric space propulsion is a technology that is used in a continuously increasing number of spacecrafts. The qualification of these propulsion systems has to run in ground-based test facilities which requires long testing times and powerful pumping systems. In these usually large test facilities, high pumping speeds are achieved with cryopumps. Cryopump operation is very expensive with respect to electrical energy and cooling water consumption. Therefore, being able to optimize pump shape, cold plate material, and pump placement in a chamber is beneficial. Pump design and tuned operating strategies can reduce costs and increase intervals between regeneration. Testing different pump configuration setups in a large facility is mostly prohibitive due to high costs and long testing times. Optimization via modelling is a better choice for design and also, later, for operation. Therefore, having a numerical model and proven guidelines at hand for optimization is very helpful. This paper describes a new model developed at DLR for the optimization of cryopump layout and operation. Model results are compared with cryopump operational and warm-up data. This validation is the basis for further optimization actions like multi-layer insulation layouts and pump cold plate upgrades, and helps in understanding and mitigating the detrimental effect of water condensates on the cryopump cold plates.

Kinetic simulations of low-pressure inductively coupled plasma: an implicit electromagnetic PIC/MCC model with the ADI-FDTD method

Low-pressure inductively coupled plasma (ICP) is promising for space electric propulsion. For the first time, an implicit electromagnetic particle-in-cell Monte Carlo collision model based on the alternating-direction-implicit finite-difference time-domain (ADI-FDTD) method is developed to investigate low-pressure xenon plasma characteristics of a miniature ICP source. The induced simulated electric field is well consistent with that calculated by the finite element method, indicating that this method can provide an accurate estimation of the electromagnetic field. The simulation time step used in the ADI-FDTD method is no longer restricted by the Courant–Friedrichs–Lewy constraints. Compared with the FDTD method, the ADI-FDTD method increases the size of the time step and significantly improves computational efficiency. The method is validated by comparing the simulated and measured electron density and plasma potential profile and reasonable agreement is reached. Therefore, the model is used to investigate the temporal and spatial distribution of plasma properties and the influence of the current amplitude of radio frequency (RF) coil, applied frequency of RF coil and neutral gas pressure on the plasma dynamics in the ionization chamber of a miniature gridded RF ion thruster. To explain the influence of the operating parameters, a concept called ‘the energy relaxation characteristics of electrons in response to the change of electric field’ is proposed and verified. The simulations also find that the oscillation frequency of plasma properties is twice the applied frequency of RF coil. The oscillation characteristics reveal the dynamic energy balance in the ICP. The experiment on the gridded RF ion thruster BHRIT-4 confirms the oscillation by measuring the plasma sheath potential.

Temperature effects on electrospray current from an externally wetted EMI-Im ionic liquid ion source

Ionic liquid ion sources are expected to be used in a wide range of applications such as space electric propulsion and focused ion beam micromachining. It is known that the backstreaming of secondary charged species generated by ion beam impacts can cause unexpected temperature rise and chemical changes in ionic liquids. This paper reports on results of heating experiments using a sharp needle emitter wetted with an ionic liquid, 1-ethyl-3-methyl imidazolium bis(trifluoromethanesulfonyl)amide, at temperatures in a range from room temperature to 120 °C. Current measurements show that positive and negative electrospray currents from the heated emitter increased as the temperature increased. Time-of-flight (TOF) mass spectrometric measurements reveal that the beam composition changed significantly with increasing temperature, indicating that charged droplets as well as ions were emitted from the heated emitter. The TOF data show that a significant fraction of the current is due to droplets at higher temperatures. On the basis of the results obtained, the size and charge of the emitted droplets are discussed. The beam is roughly estimated to contain charged droplets with a diameter of around 20 nm at 120 °C.

Experimental research on micro thrust measurement of ion thruster under variable working conditions

In view of the application requirements for accurately measuring the micro thrust of ion thrusters in space electric propulsion, an experimental study of micro thrust measurement under variable working conditions was carried out based on an 8 cm ion thruster. By combining highly sensitive thrust response parameters such as gas flow, anode voltage and grid voltage, the ion thruster can achieve high-precision continuous thrust adjustment in the range of 1–18.6 mN. The calibration results of the thrust measuring device were compared in the cold state and the stable state of thermal equilibrium of the ion thruster, respectively, and the measured thrust was compared with the thrust value calculated based on beam current characteristic measurement. The experimental results showed that compared with the cold calibration, the calibration results of the thrust measuring device in the stable state of thermal equilibrium of the ion thruster are more suitable for measuring the thrust of the ion thruster, and the measured thrust value is in good agreement with the thrust value calculated based on beam current characteristic measurement, and the relative error of the two within the research range is less than 4%.

Ammonia fuelled space electric propulsion systems using C12A7:e- electride as electron emitter

Given the significant changes taking place in the geopolitical global situation, and the derived supply chain issues for some traditional electric propulsion propellants like xenon, alternative propellants issue is perceived as a strategic topic to tackle, and ammonia (NH3) is becoming one serious candidate. Ammonia is increasingly being investigated to extend green hydrogen use by overcoming the storage and transportation issues of hydrogen. Ammonia characteristics like its superior energy density and low temperature and/or pressure needs for storage (10 bar at 20°C for liquid ammonia), make it very valuable for simplified unexpensive energy storage and transportation, and these characteristics makes it also especially suitable for on-board spacecrafts electric propulsion purposes. Based on our research activities on ammonia generation and dissociation processes with C12A7:e- electride as catalyst, this work will describe the most relevant characteristics and properties associated to the advantages of using the cheap and abundant ammonia, and will also present and discuss the results of the first successful tests performed with ammonia as fuel for a C12A7:e- electride based neutralizer, including relevant endurance tests in operation conditions. Additionally, the dual application as propellant and on-board energy generation system of ammonia will also be discussed.

Novel concept of tailorable magnetic field and electron pressure distribution in a magnetic nozzle for effective space propulsion

Magnetic nozzle appears to be a practical solution for prolonged space missions. For efficient handling of the spaceship, an in-flight solution to customize the thrust from the magnetic nozzle is essential. Here a new concept of three-thick coils system is proposed for tailoring the magnetic field in-flight in accordance with electron pressure distribution. The role of peak position of the pressure and its axial gradient is also uncovered for realizing higher thrust. About three-fold increase in thrust is observed when the electron temperature is raised to ∼2.5 times of its original value at the exit plane. The set-up is optimized for its best performance and efficient use in the electric space propulsion sector with thrust approaching 5 mN. In particular, this can contribute to the attitude control or the precision pointing of the spacecraft, the technology for removal of space debris and manipulating the ion momentum flux lost to a wall or unsteady laser produced plasma flow in a magnetic nozzle.

Ground-Based Experiment for Electric Propulsion Thruster Plume—Magnetic Field Interaction

Electric space propulsion is a technology which is employed on a continuously increasing number of spacecrafts. While the current focus of their application area is on telecommunication satellites and on space exploration missions, several new ideas are now discussed that go even further and apply the thruster plume particle flow for transferring momentum to targets such as space debris objects or even asteroids. In these potential scenarios, the thruster beam impacts on distant objects and subsequently generates changes in their flight path. One aspect which so far has not been systematically investigated is the interaction of the charged particles in the propulsion beam with magnetic fields which are present in space. This interaction may result in a deflection of the particle flow and consequently affect the aiming strategy. In the present article, basic considerations related to the interaction between electric propulsion thruster plumes and magnetic fields are presented. Experiments with respect to these questions were conducted in the high-vacuum plume test facility for electric thrusters (STG-ET) of the German Aerospace Center in Göttingen utilizing a gridded ion thruster, an RIT10/37, and a Helmholtz coil to generate magnetic fields of varying field strength. It was possible to detect a beam deflection on the RIT ion beam caused by a magnetic field with an Earth-like magnetic field strength.

Diversity of Physical Processes: Challenges and Opportunities for Space Electric Propulsion

The renewed interest in space exploration has led to the growth in research efforts pertaining to advanced space propulsion systems, including highly efficient electric propulsion systems. Although already tested in space many decades ago and being currently employed on various space platforms and thousands of satellites, these systems are yet to reach their full potential for applications on orbit and in deep space. One specific feature of space electric propulsion is the large diversity of physical processes used in this technology, which is not typical for many other types of propulsion systems used in transport, such as those used by airplanes or automobiles. Various physical processes and mechanisms underpin different electric propulsion technologies and should be integrated to drive the future science and technology of space electric propulsion systems. This opinion article briefly highlights this feature of space electric propulsion and outlines some challenges and opportunities that follow from this diversity.

Nano-graphite field-emission cathode for space electric propulsion systems

Improving the thruster efficiency is a crucial challenge for the development of space electric propulsion systems, especially advanced air-breathing thrusters utilizing the surrounding rarefied atmosphere as fuel. A significant reduction in thruster power consumption can be achieved by using field emission (FE) cathodes that do not require heating and have the highest energy efficiency. In this work, we study FE from nano-graphite thin films, consisting of carbon nanostructures with a high aspect ratio, and demonstrate their suitability for use in the space electric propulsion systems. The films shown appropriate FE characteristics in a wide range of gas pressures at high current loads in constant and pulsed operation modes. Based on the obtained experimental results, nano-graphite cathodes were employed for the design of an electron gun with increased reliability and minimized energy losses associated with electron extraction. The possibility of using such a gun in a specific air-breathing satellite operating in low Earth orbits is demonstrated.

Single- and two-photon absorption laser-induced fluorescence spectroscopy in rare gases for gridded ion thruster diagnostics

Methods based on laser-induced fluorescence spectroscopy are widely used for spatially resolved non-intrusive diagnostics of atomic or molecular densities and velocity distributions in plasma applications. With regard to electric space propulsion, one focus is on the investigation of rare gases such as xenon or krypton, which are currently the favored propellants in gridded ion- and Hall-effect thrusters. For gridded ion engines, diagnostics of neutral atoms is of interest since charge-exchange processes between neutrals and ions are the main driver of accelerator grid erosion, which limits the lifetime of a gridded ion thruster. Extending the capabilities of the advanced electric propulsion diagnostics platform which has been developed by the IOM and partners, single- and two-photon absorption laser-induced fluorescence diagnostics have been set-up recently at our institute. Both experimental set-ups, and as a series of first applications, measurements of krypton neutrals in the plume of the radiofrequency ion thruster RIT-10 (ArianeGroup GmbH), and xenon neutrals within the discharge chamber of a gridded radiofrequency ion source developed at IOM, are presented.

Performance comparison of LaB6 and C12A7:e-emitters for space electric propulsion cathodes

This work, based on an EU-funded project, NEMESIS, is aiming at developing electride-based cathode technology which is compatible with all kinds of electric propulsion systems requiring neutralization. Its target is to demonstrate and validate the performance of a novel C12A7:e-electride material as electron emitter instead of traditional thermionic emitters such as lanthanum hexaboride, LaBe, or barium oxide, BaO. In this study, a fair comparison between LaBe and C12A7:e-samples was performed both addressing pure material characterization parameters as well as comparing performance as cathodes under different architectures and operational conditions. In this case, a current/cathode power ratio around 3 mA/W was obtained when using the C12A7:e-sample in a plasma environment with Ar, which is approximately one order of magnitude higher compared to the LaB6 sample.

Evolution in maximised efficiency engines for space applications: from GIESEPP to GIESEPP MP

Evolution in maximised efficiency engines for space applications: from GIESEPP to GIESEPP MP The leading European companies for orbital space propulsion together with satellite integrators and in EP (electric propulsion) specialised companies, universities and research partners have joined efforts to boost the European electric space propulsion capabilities in the field of efficiency optimised gridded ion engine technology.

First Helicon Plasma Physics and Applications Workshop

The First Helicon Plasma Physics and Applications Workshop was held on September 23−24, 2021, through Zoom Cloud Meeting, instead of in an on-site gathering, due to the COVID-19 pandemic. It was convened by Rod Boswell (IOC) and Guangnan Luo (LOC), and organised by Lei Chang’s group. The workshop attracted 110 registrations and ∼100 online audiences from ∼30 affiliations. There were 33 presentations covering the various fundamental physics of helicon plasma and its applications to space electric propulsion, material processing, and magnetic confinement fusion. This paper highlights the presentations, discussions, and perspectives given in the workshop, serving as reference for the helicon community.

Thrust measurement and thrust balance development at DLR\u2019s electric propulsion test facility

Electric space propulsion thrusters only produce low thrust forces. For the fulfillment of a space mission this implies long thruster runtimes, and this entails long qualification times on ground. For such long testing times, a ground facility requires a vacuum chamber and a powerful pumping system which can guarantee high vacuum over extended times and under thruster gas load. DLR’s STG-ET is such a ground test facility. It has a high pumping capability for the noble gases typically used as propellants. One basic diagnostic tool is a thrust measurement device, among various other diagnostic systems required for electric propulsion testing, e.g. beam diagnostics. At DLR we operate a thrust balance developed by the company AST with a thrust measurement range of 250 mN and capable of thruster weights up to 40 kg. Adversely, it is a bulky and heavy device and all upgrades and qualification work needs to be done in a large vacuum chamber. In order to have a smaller device at hand a second thrust stand is under development at DLR. The idea is to have a light and compact balance that could also be placed in one of the smaller DLR vacuum chambers. Furthermore, the calibration is more robust and the whole device is equipped with a watercooled housing. First tests are promising and showed a resolution well below 1 mN. In this paper we give background information about the chamber, describe the basics of thrust measurement and the development of a new balance.