Field Emission Electric Propulsion Thrusters Research Articles

Simulation of liquid cone formation on the tip apex of indium field emission electric propulsion thrusters

Field emission electric propulsion (FEEP) thrusters possess excellent characteristics, such as high specific impulse, low power requirements, compact size and precise pointing capabilities, making them ideal propulsion devices for micro-nano satellites. However, the detection of certain aspects, such as the evolution process of the liquid cone and the physical quantities at the cone apex, proves challenging due to the minute size of the needle tip and the vacuum environment in which they operate. Consequently, this paper introduces a computational fluid dynamics (CFD) model to gain insight into the formation process of the liquid cone on the tip apex of indium FEEP. The CFD model is based on electrohydrodynamic (EHD) equations and the volume of fluid (VOF) method. The entire cone formation process can be divided into three stages, and the time-dependent characteristics of the physical quantities at the cone apex are investigated. The influences of film thickness, apex radius size and applied voltage are compared. The results indicate a gradual increase in the values of electrostatic stress and surface tension stress at the cone apex over an initial period, followed by a rapid escalation within a short duration. Apex configurations featuring a small radius, thick film and high voltage exhibit a propensity for liquid cone formation, and the cone growth time decreases as the film thickness increases. Moreover, some unstable behavior is observed during the cone formation process.

On the cone-to-jet transition region and its significance in electrospray propulsion

A novel method for identifying the cone-to-jet transition region is proposed to characterize the electrospray dynamics, and thus the Taylor cone, a fundamental principle of ion extraction from the tip of the Taylor cone in an electric field to produce μN thrust for nanosatellites' application. A numerical model is developed and validated against the droplet's diameter with a proper characterization of the cone-to-jet transition region based on the hydrodynamic pressure gradient at the cone-jet axis. The results show that the cone-to-jet transition region is reduced when the electric potential increases and the flow rate decreases. Moreover, the current density is found to be higher for the lower flow rate and large electric potential. Interestingly, the current density at the jet axis increases significantly at the flow rate of 1 cc/hr. This implies that below a certain droplet diameter, the current density carried by the jet interface is completely migrated into the jet, and the substantial change in the current density occurs within the cone-to-jet transition region. When comparing the current density along the jet axis, the small flow rate and large electric potential carry the maximum current density at the jet interface. The identification of cone-to-jet transition region is central in modeling the ion beam trajectory via initial current density profile required for particle-in-cell calculation for assuring the performance of a field emission electric propulsion thruster.

IFM Nano Thruster performance studied by experiments and numerical simulations

Field emission electric propulsion thrusters are characterised by their low thrust range, which makes them ideal for the precise control of spacecraft. Decisive for precise control are the properties of the thruster ion beam, which includes the beam divergence angle and the thrust vector. The analysis of these properties is also necessary in order to be able to estimate the interactions of the beam with components of the spacecraft. Due to such interactions, solar panels or electrical instruments on board the spacecraft could be damaged by sputtering effects.The spatial ion current density and energy distribution of a test crown emitter beam, with different specifications compared to the IFM Nano thruster, were examined experimentally with a diagnostic system, including Faraday cups and a retarding potential analyser. In addition to the analysis of the beam profile of an emitting crown, a single emitting needle was analysed. Based on these experimental analyses, an ion trajectory simulation model was developed to determine the theoretical ion current density distribution. This model includes the properties of a liquid metal ion source, where the ion trajectories start from their point of origin, the so-called Taylor cone jet cap. The benchmark of the model shows that the thrust vector and divergence angle correspond to the experimental results and shows the identical calculations for different thruster parameters, like emission current and electrode voltage. The simulation allows for the optimisation of existing and novel thruster geometries in terms of performance, reliability and longevity.

Metal Plasma Thruster for Small Satellites

This paper describes a metal plasma thruster (MPT) that is suited to nano- and microsatellites. The MPT uses solid metal propellant (hence requires no liquids, gases, flow valves, or flow controls), has no moving parts, generates approximately , and may be packaged into , 1-U, and multiple-U increments. A vacuum arc generates a metal plasma plume that accelerates into space at (platinum, Pt) to (magnesium, Mg) to generate thrust (). The range of specific impulse is from 693 s (Pt) up to 3119 s (Mg). Spacecraft interface is simple: the autonomous MPT requires only a DC input (7–28 V) and a start/stop control pulse from the satellite. The MPT is simpler and more rugged than ion engines and Hall thrusters, which are not easily scaled down in mass and power for satellites, and provides superior total impulse per kilogram. The MPT is easier to operate than equivalent-sized alternatives: field emission electric propulsion thruster, xenon gas radiofrequency thruster, and pulsed plasma thruster.

Characteristics of liquid film on the tip surface of indium field-emission electric propulsion thrusters

The characteristics of the propellant film on the tip surface of the field-emission electric propulsion thruster have a close relationship with the work state of the thruster. In this paper, the distributions of both the liquid film thickness and the fluid velocity along the needle tip have been calculated by building the liquid mechanics equations and a cone coordinate system suitable for the tip. The results show that the film thickness is tens to hundreds of nanometers and the fluid velocity along the needle tip is from tens of µm/s to a few mm/s.

Polyatomic ions from a high current ion implanter driven by a liquid metal ion source.

High current liquid metal ion sources are well known and found their first application as field emission electric propulsion thrusters in space technology. The aim of this work is the adaption of such kind of sources in broad ion beam technology. Surface patterning based on self-organized nano-structures on, e.g., semiconductor materials formed by heavy mono- or polyatomic ion irradiation from liquid metal (alloy) ion sources (LMAISs) is a very promising technique. LMAISs are nearly the only type of sources delivering polyatomic ions from about half of the periodic table elements. To overcome the lack of only very small treated areas by applying a focused ion beam equipped with such sources, the technology taken from space propulsion systems was transferred into a large single-end ion implanter. The main component is an ion beam injector based on high current LMAISs combined with suited ion optics allocating ion currents in the μA range in a nearly parallel beam of a few mm in diameter. Different types of LMAIS (needle, porous emitter, and capillary) are presented and characterized. The ion beam injector design is specified as well as the implementation of this module into a 200 kV high current ion implanter operating at the HZDR Ion Beam Center. Finally, the obtained results of large area surface modification of Ge using polyatomic Bi2+ ions at room temperature from a GaBi capillary LMAIS will be presented and discussed.

The beam divergence of an indium LMIS at a distance of 50μm as determined by plasma diagnostic measurements

The current-dependent beam divergence at a distance of 50 μm from an indium-liquid metal ion source is derived from experimental data obtained by measuring the beam spread with a 3D Plasma diagnostic system at a distance of 10 cm from the needle tip. The observed relationship between emission current and beam divergence in vicinity of the emitting needle is used to design a focusing electrode for a field-emission electric propulsion thruster operating at currents up to 150 μA. Another application involves focused ion beam columns which may choose to forego a beam-limiting aperture, such as LMIS-based rapid machining tools with large beam currents.

Development of Electric and Chemical Microthrusters

The increasing application of microsatellites (from 10 kg up to 100 kg) as well as CubeSats for a rising number of various missions demands the development of miniaturized propulsion systems. Fotec and The University of Applied Sciences at Wiener Neustadt is developing a number of micropropulsion technologies including both electric and chemical thrusters targeting high performance at small scales. Our electric propulsion developments include a series of FEEP (field emission electric propulsion) thrusters, of which the thrust ranges fromμN to mN level. The thrusters are highly integrated into clusters of indium liquid-metal-ion sources that can provide ultralow thrust noise and long-term stability. We are also developing a micro PPT thruster that enables pointing capabilities for CubeSats. For chemical thrusters, we are developing novel micromonopropellant thrusters with several hundred mN as well as a 1–3 N bipropellant microrocket engine using green propellants and high specific impulse performance. This paper will give an overview of our micropropulsion developments at Fotec, highlighting performance as well as possible applications.

ORBIT MAINTENANCE FOR SMALL SATELLITES USING ELECTRIC PROPULSION

ABSTRACTControlling a small satellite position using low thrust electric propulsion requires continues control. Continues control means that the orbit control system will command the thrusters every small period of time so that the error will not accumulate. This is due to the fact that low thrust level thrusters are unable to correct an accumulated — relatively — big error. In this paper a control algorithm is developed for controlling a small satellite with Field Emission Electric Propulsion (FEEP) thruster. This algorithm takes into account the errors induced in the GPS receiver and an available thrust level of 1 mN. in the satellite velocity direction only Controller gains are chosen so as to optimize the system equivalent noise bandwidth. This approach is based on estimating the states based on the measurements and controlling the satellite position using the estimated states.This algorithm uses the modulation capability of the thruster to produce a proportional thrust for disturbances compensation.A state estimation technique is used to estimate satellite position from GPS receiver. Batched least square estimator is used. The estimated states are used with a PO discrete controller to correct the satellite position. The results pointed out the potential of using state estimation and showed that the satellite position could be kept in orbit with an error in altitude less than one meter using a single 1 mN FEEP thruster.

ORBIT MAINTENANCE FOR SMALL SATELLITES USING ELECTRIC PROPULSION

Controlling a small satellite position using low thrust electric propulsion requires continues control. Continues control means that the orbit control system will command the thrusters every small period of time so that the error will not accumulate. This is due to the fact that low thrust level thrusters are unable to correct an accumulated — relatively — big error. In this paper a control algorithm is developed for controlling a small satellite with Field Emission Electric Propulsion (FEEP) thruster. This algorithm takes into account the errors induced in the GPS receiver and an available thrust level of 1 mN. in the satellite velocity direction only Controller gains are chosen so as to optimize the system equivalent noise bandwidth. This approach is based on estimating the states based on the measurements and controlling the satellite position using the estimated states. This algorithm uses the modulation capability of the thruster to produce a proportional thrust for disturbances compensation. A state estimation technique is used to estimate satellite position from GPS receiver. Batched least square estimator is used. The estimated states are used with a PO discrete controller to correct the satellite position. The results pointed out the potential of using state estimation and showed that the satellite position could be kept in orbit with an error in altitude less than one meter using a single 1 mN FEEP thruster.

Numerical Model of Cesium Ion Beams in Field Emission Electric Propulsion Thrusters

A numerical model of the ion beam in field emission electric propulsion thrusters is presented. The objective of the study is to develop a computational tool capable of predicting the effects of thruster configuration parameters (geometry, voltage distribution between the electrodes) on thruster performance (beam divergence, ion exhaust velocity). A two-dimensional, steady model has been adopted. A simple emission model has been assumed, neglecting the detailed, microscopic description of the physical emission mechanism. The electric field is evaluated by a finite element scheme; boundary conditions are provided by an auxiliary technique based on the placement of fictitious charges in proximity to the electrode surfaces. The beam is simulated through a discrete number of charged particles; the effects of the associated space charge density distribution on the electric field is considered. The results of the simulations of typical thruster configurations are presented and discussed. The need to take space charge effects into account to attain a sufficiently accurate solution is indicated: the beam charge density modifies the electric potential distribution generated by the electrodes voltage difference, causing a decrease in the local electric field at emitter tip and an increase in beam divergence. Neutralization at a finite distance must be taken into account to fully appreciate the effects of different voltage distributions between the electrodes: a higher accelerator electrode voltage determines a lower ion exhaust velocity and a larger beam divergence. The capability of an additional, decelerating neutral electrode to focus the ion beam is also investigated, showing that only little improvement in beam containment is attained.

A design study of a microthrust balance for space applications

Spacecraft drift from their desired orbit as a result of being subjected to external forces, such as gravity, atmospheric drag or solar pressure. In situ propulsion systems are employed to maintain a spacecraft in a specific orbit. Field emission electric propulsion (FEEP) has been proposed as a technique for maintaining orbits to very high precision for applications such as the measurement of the effects of gravitational waves. The specification of the FEEP thrusters is very demanding and they require calibration over a range of 1- to an accuracy of in the frequency range of 0.1 to 10 Hz. A design study has been completed of two balances, one with a real pivot and the other with a virtual pivot, for calibrating FEEP thrusters.

Pulsed thrust measurements using laser interferometry

An optical interferometric proximeter system (IPS) for measuring thrust and impulse bit of pulsed electric thrusters was developed. Unlike existing thrust stands, the IPS-based thrust stand offers the advantage of a single system that can yield electromagnetic interference-free, high accuracy (<2% error) thrust measurements within a very wide range of impulses (100 μN s to above 10 N s) covering the impulse range of all known pulsed plasma thrusters. In addition to pulsed thrusters, the IPS is theoretically shown to be capable of measuring steady-state thrust values as low as 20 μN for microthrusters such as the field emission electric propulsion thruster. The IPS-based thrust stand relies on measuring the dynamic response of a swinging arm using a two-sensor laser interferometer with 10 nm position accuracy. The wide application of the thrust stand is demonstrated with thrust measurements of an ablative pulsed plasma thruster and a quasi-steady magnetoplasmadynamic thruster.

`Galileo Galilei' flight experiment on the equivalence principle with field emission electric propulsion

An experiment to test the equivalence of inertial to gravitational (passive) mass in space offers two main advantages: a signal about a factor of a thousand larger than on Earth and the possibility of exploiting the absence of weight. `Galileo Galilei' (GG) is a small satellite mission currently under study in Italy with the financial support of ASI (Agenzia Spaziale Italiana). The mission concerns a small, low Earth satellite ( total mass, altitude) with two objectives. One is scientific, in the field of fundamental physics, and the other technological within the framework of spacecraft propulsion and drag compensation. The scientific goal is to test the equivalence principle to one part in , four orders of magnitude better than the best ground results. The technological goal is a full, comprehensive test of FEEP (field emission electric propulsion) thrusters for accurate drag compensation, a technology developed in Europe by the ESA (European Space Agency) which is likely to become an essential component of all space experiments which require measurement of small forces. The GG experiment is based on novel concepts and does not require low temperatures.