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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. [ABSTRACT FROM AUTHOR]

The neutralizer-free ion thruster based on radio frequency (RF) acceleration was proposed as an attractive approach. This comment presents problems arising from such RF acceleration. Two mayor limitations in developing the neutralizer-free RF ion thruster into a practical technology are found to be involved with the grounded grid: (a) the high frequency bias required by low ion impingement from the upstream; (b) serious backstreaming of electron due to the grounded grid that cannot establish a potential barrier. Limitation (a) will degrade the ion production efficiency and cause more serious thermal deformation of the biased grids, especially for large-sized RF ion thrusters. Limitation (b) is the Achilles' heel of a neutralizer-free RF-biased ion thruster: it in fact operates from the very beginning in the failure mode occurring in the final stage of seriously-eroded DC (direct current) biased ion thruster. The backstreaming electrons will substantially reduce the lifetime of neutralizer-free RF ion thrusters. In the current stage, the RF acceleration is not a promising neutralizer-free technology. [ABSTRACT FROM AUTHOR]

The development of plasma-based propulsion thrusters for spacecraft has seen a rapid growth over the past few decades, with the number of spacecraft including small satellites and Cubesats increasing exponentially. Although traditional chemical propulsion is still widely employed in space flights, it cannot meet the more challenging requirements for deep space travel due to low specific impulse. Electric propulsion thrusters have already helped humans travel further from Earth and have the potential to be developed for interstellar flights due to their advantages such as high velocity increments, long operational lifetimes, high impulse-to-weight ratios and high impulse-to-power ratios. The electrodynamic thrusters have significant potential for applications in the remote regions of space, and several types of electrodynamic plasma thrusters are currently under investigation. In this paper we present conceptual experiments to study a miniaturized Rotamak-type device initially proposed for the controlled thermonuclear fusion, with a view to assess its potential for the application as a small space thruster. An outline of the physical characteristics of the experiments that has been carried out, and measurements were done to try to elucidate the important mechanisms at work in the Rotamak, which will help design next-generation thruster capabilities. A discussion is also presented about the Rotamak systems and the opportunities they present for space applications. [ABSTRACT FROM AUTHOR]

The discrete control of ion energy and flux is of increasing importance to industrially relevant plasma sources. The ion energy distribution functions (IEDFs) and net ion flux incident upon material surfaces in intermediate pressure (≈133 Pa, 1 Torr) radio-frequency capacitively coupled plasmas (rf CCPs) are coupled to the spatio-temporal sheath dynamics and resulting phase-averaged sheath potential. For single frequency driven discharges this co-dependence of ion energy and flux on the sheath potential limits the range of accessible operating regimes. In this work, experimentally benchmarked 2D fluid/Monte-Carlo simulations are employed to demonstrate quasi-independent control of the ion flux and IEDF incident upon plasma facing surfaces in a collisional (≈200 Pa, 1.5 Torr argon) rf hollow cathode discharge driven by multi-harmonic (n ⩾ 2) tailored voltage waveforms. The application of variable phase offset n = 5 tailored voltage waveforms affords a significant degree of control over the ion flux and mean ion energy , modulating each by factors of 2.9 and 1.6, respectively as compared to 1.8 and 1.6, achieved via n = 2 dual-frequency voltage waveforms. The disparate modulations achieved employing n = 5 tailored voltage waveforms demonstrate a significant degree of independent control over the mean ion energy and ion flux for collisional conditions, enabling access to a wider range of operational regimes. Maximising the extent to which ion energy and flux may be independently controlled enables improvements to plasma sources for technological applications such as plasma assisted material manufacture and spacecraft propulsion. [ABSTRACT FROM AUTHOR]

With the increased interest in electric propulsion for space applications, a wide variety of electric thrusters have emerged. For many years, Hall effect thrusters have been the selected technology to sustain observation and telecommunication satellites thanks to their advantageous service lifetime, their high specific impulse and high power to thrust ratio. Despite several studies on the topic, the Hall thruster electric discharge remains still poorly understood. With the increase of available computing resources, numerical simulation becomes an interesting tool in order to explain some complex plasma phenomena. In this paper, a fluid model for plasma flows is presented for the numerical simulation of space thrusters. Fluid solvers often exhibit strong hypotheses on electron dynamics via the drift-diffusion approximation. Some of them use a quasi-neutral assumption for the electric field which is not adapted near walls due to the presence of sheaths. In the present model, all these simplifications are removed and the full set of plasma equations is considered for the simulation of low-temperature plasma flows inside a Hall thruster chamber. This model is implemented in the unstructured industrial solver AVIP, efficient on large clusters and adapted to complex geometries. Electrical sheaths are taken into account as well as magnetic field and majors collision processes. A particular attention is paid on a precise expression of the different source terms for elastic an inelastic processes. The whole system of equations with adapted boundary conditions is challenged with a simulation of a realistic 2D r–z Hall thruster configuration. The full-fluid simulation exhibits a correct behavior of plasma characteristics inside a Hall effect thruster. Comparisons with results from the literature exhibit a good ability of AVIP to model the plasma inside the ionization chamber. Finally a specific attention was brought to the analysis of the thruster performances. [ABSTRACT FROM AUTHOR]

In this paper, we explore from a conceptual standpoint the possibility of using natural astrophysical sources to accelerate spacecraft to relativistic speeds. We focus on light sails and electric sails, which are reliant on momentum transfer from photons and protons, respectively, because these two classes of spacecraft are not required to carry fuel on board. The payload is assumed to be stationed near the astrophysical source, and the sail is subsequently unfolded and activated when the source is functional. By considering a number of astrophysical objects such as massive stars, microquasars, supernovae, pulsar wind nebulae, and active galactic nuclei, we show that terminal speeds approaching the speed of light might be realizable under idealized circumstances provided that sufficiently advanced sail materials and control techniques exist. We also investigate the constraints arising from the sail's material properties, the voyage through the ambient source environment, and the passage through the interstellar medium. While all of these considerations pose significant challenges to spacecraft, our analysis indicates that they are not insurmountable in optimal conditions. Finally, we sketch the implications for carrying out future technosignature searches. [ABSTRACT FROM AUTHOR]

High-temperature superconducting (HTS) magnetic bearings have great potential in the field of rotating machinery, such as flywheels and momentum wheels, because of their low friction and self-stabilizing capability. The study of rotational loss in HTS bearing is of great significance to achieve better efficiency. Previous research indicates that low rotational loss has a close relation with hysteresis loss in HTSs and eddy-current losses in the surrounding metal materials. A low coefficient of friction (<10−8) at low speed is attractive for high precision instruments. This paper first presents a short overview of low rotational loss in HTS bearings, and then a special design of HTS bearing is given for application in micro-thrust measurements of space propulsion. The load capacity of the bearing allows for up to 10 kg, which means that the entire thruster system can be suspended to avoid measuring the internal force. Two different measuring methods are designed to fit several operational modes of the micro-thruster, and the measured thrust level is from 1 mN to 100 mN. The bearing system shows great response to short pulse modes of the micro-thruster, and an impulse of 1 mN·s was identified accurately. The measured thrust shows great agreement with the designed values of the micro-thruster, indicating the effective application of the HTS bearing for micro-thrust measurements. [ABSTRACT FROM AUTHOR]

Specific energy (i.e. energy per unit mass) is one of the most fundamental and consequential properties of a fuel source. In this work, a systematic study of measured fusion cross-sections is performed to determine which reactions are potentially feasible and identify the fuel cycle that maximizes specific energy. This reveals that, by using normal hydrogen to breed deuterium via neutron capture, the conventional catalyzed D–D fusion fuel cycle can attain a specific energy greater than anything else. Simply surrounding a catalyzed D–D reactor with water enables deuterium fuel, the dominant stockpile of energy on Earth, to produce as much as 65% more energy. Lastly, the impact on space propulsion is considered, revealing that an effective exhaust velocity exceeding that of deuterium-helium-3 is theoretically possible. [ABSTRACT FROM AUTHOR]

A direct measurement torsional thrust balance for electric thrusters has been developed to test pulsed plasma thrusters (PPTs) and other electric thrusters. The current configuration is able to measure impulse bits between 14 and 79 μN · s with a resolution of 0.05 to 0.10 μN · s. The direct measurement technique, in which the thruster is mounted on the thrust balance, is considered the most suitable and accurate for applications with PPTs. The use of flexural pivots with low stiffness makes the balance compact, with a smaller torsional arm. An electrostatic calibration device (ECD) was designed and built to calibrate the thrust balance, simulate the thrust of PPTs, and to allow forthcoming tests of the double discharge pulsed plasma thrusters developed at the Associated Laboratory of Combustion and Propulsion of the Brazilian National Institute for Space Research at the Electric Space Propulsion Laboratory. The maximum impulse bit measurement error achieved was as low as 10%, lower than the average of 12%–15%, typical for this type of thrust balance. The ECD was calibrated using a previously calibrated precision analytic scale. The ECD then can carry this calibration traceability to the thrust balance, providing a higher level of accuracy. [ABSTRACT FROM AUTHOR]

The quiet Sun (QS) and helmet streamer (STR) are generally considered to account for sources of slow solar wind with low Alfvénicity and low proton temperature. The solar wind with high Alfvénicity is often associated with coronal holes (CHs). Recently, the solar wind measured by the Advanced Composition Explorer spacecraft at 1 au was mapped back to its magnetic footpoints. Depending on the proximity of the solar wind footpoints to a given coronal or heliospheric structure, we classify solar winds into different types based on their sources: QS, STR, active region (AR), and CH. Here, we compare the properties of the solar winds originated from QS, STR, AR, and CH using 2 hr data. We find that at solar maximum 34% of the quiet-Sun-associated slow wind (VSW < 450 km s−1) has high Alfvénicity (). This significantly higher proportion of Alfvénic fluctuations indicates that the quiet-Sun-associated wind at 1 au has similar properties as fast wind, which originates from the CH. Accordingly, we speculate that this type of solar wind at 1 au could come from open fields within the quiet-Sun region. This observational study will help us understand more about the coronal source regions of the solar wind in interplanetary space. [ABSTRACT FROM AUTHOR]

An axisymmetric, finite difference frequency domain model is used to study the wave propagation and power absorption in a helicon plasma thruster operating inside a laboratory vacuum chamber. The magnetic field is not purely axial and the plasma beam is cylindrical in the source and divergent in the magnetic nozzle. The influence of the magnetic field strength, plasma density, electron collision frequency and geometry on the wavefields and the power absorption maps is investigated, showing different power deposition patterns. The electromagnetic radiation is not confined to the source region but propagates into the nozzle divergent region, and indeed the power absorption there is not negligible. For the impedance at the antenna, the reactance is rather constant but the resistance is very dependent on operation parameters; optimal parameter values maximizing the resistance are found. [ABSTRACT FROM AUTHOR]

This article numerically investigates a micro-thruster for small satellites which utilizes plasma actuators to heat and accelerate the flow in a micro-channel with rarefied gas in the slip flow regime. The inlet plenum condition is considered at 1 Torr with flow discharging to near vacuum conditions (<0.05 Torr). The Knudsen numbers at the inlet and exit planes are ~0.01 and ~0.1, respectively. Although several studies have been performed in micro-hallow cathode discharges at constant pressure, to our knowledge, an integrated study of the glow discharge physics and resulting fluid flow of a plasma thruster under these low pressure and low Knudsen number conditions is yet to be reported. Numerical simulations of the charge distribution due to gas ionization processes and the resulting rarefied gas flow are performed using an in-house code. The mass flow rate, thrust, specific impulse, power consumption and the thrust effectiveness of the thruster are predicted based on these results. The ionized gas is modelled using local mean energy approximation. An electrically induced body force and a thermal heating source are calculated based on the space separated charge distribution and the ion Joule heating, respectively. The rarefied gas flow with these electric force and heating source is modelled using density-based compressible flow equations with slip flow boundary conditions. The results show that a significant improvement of specific impulse can be achieved over highly optimized cold gas thrusters using the same propellant. [ABSTRACT FROM AUTHOR]

Microdischarges are an important type of plasma discharge that possess several unique characteristics, such as the presence of a stable glow discharge, high plasma density and intense excimer radiation, leading to several potential applications. The intense and controllable gas heating within the extremely small dimensions of microdischarges has been exploited in micro-thruster technologies by incorporating a micro-nozzle to generate the thrust. This kind of micro-thruster has a significantly improved specific impulse performance compared to conventional cold gas thrusters, and can meet the requirements arising from the emerging development and application of micro-spacecraft. In this paper, we performed a self-consistent 2D particle-in-cell simulation, with a Monte Carlo collision model, of a microdischarge operating in a prototype micro-plasma thruster with a hollow cylinder geometry and a divergent micro-nozzle. The model takes into account the thermionic electron emission including the Schottky effect, the secondary electron emission due to cathode bombardment by the plasma ions, several different collision processes, and a non-uniform argon background gas density in the cathode–anode gap. Results in the high-pressure (several hundreds of Torr), high-current (mA) operating regime showing the behavior of the plasma density, potential distribution, and energy flux towards the hollow cathode and anode are presented and discussed. In addition, the results of simulations showing the effect of different argon gas pressures, cathode material work function and discharge voltage on the operation of the microdischarge thruster are presented. Our calculated properties are compared with experimental data under similar conditions and qualitative and quantitative agreements are reached. [ABSTRACT FROM AUTHOR]

Electric space propulsion is an intensively developing field addressing new demands and challenges for long-term spacecraft operation. Many novel plasma propulsion concepts aim to find new acceleration principles, use alternative propellants, upscale or downscale thrusters for large thrust or for very small spacecrafts etc. In this work we review the neutralizer-free concepts, where both positive and negative particles are extracted and accelerated from plasmas. We can divide these concepts into three main categories, defined by their acceleration principle: (i) neutral beam generation, (ii) plasma acceleration/expansion and (iii) bipolar beam acceleration. We describe the basic physical principles and evaluate the main advantages and drawbacks in view of general space applications. We also present here further detail on a recent concept where RF voltages are used to accelerate quasi-simultaneously positive ions and electrons from the same source. [ABSTRACT FROM AUTHOR]

Ionic liquid ion sources (ILIS) have a broad range of applications ranging from focused ion beams (FIB) to spacecraft propulsion. For space thrusters, having a beam of ions of high velocity and low energy spread results in high specific impulse (a measure of the efficiency of propellant consumption) and high power efficiency. For FIB applications, these traits benefit focusing and resolution respectively. Since typical current levels emitted by individual ILIS range from tens of nA to a few μA, both applications often require using arrays of emitters in order to increase the total beam current; i.e. the thrust level in the case of a thruster or the throughput in ion milling applications. So far, most studies of emitter arrays have dealt with experimental measurements of global, spatially-averaged, properties and theoretical analyses generally assume that all emitters operate in a similar manner, so that scaling of the current is linear with the number of emitters. In this work we use a current probe to explore the uniformity properties of the current emitted by an array of porous emitters. We present measurements of the current density of both the array and individual tips, and analyze the reason and implications of the experimental observations. [ABSTRACT FROM AUTHOR]

Propulsion is required for satellite motion in outer space. The displacement of a satellite in space, orbit transfer and its attitude control are the task of space propulsion, which is carried out by rocket engines. Electric propulsion uses electric energy to energize or accelerate the propellant. The electric propulsion, which uses electrical energy to accelerate propellant in the form of plasma, is known as plasma propulsion. Plasma propulsion utilizes the electric energy to first, ionize the propellant and then, deliver energy to the resulting plasma leading to plasma acceleration. Many types of plasma thrusters have been developed over last 50 years. The variety of these devices can be divided into three main categories dependent on the mechanism of acceleration: (i) electrothermal, (ii) electrostatic and (iii) electromagnetic. Recent trends in space exploration associate with the paradigm shift towards small and efficient satellites, or micro- and nano-satellites. A particular example of microthruster considered in this paper is the micro-cathode arc thruster (µCAT). The µCAT is based on vacuum arc discharge. Thrust is produced when the arc discharge erodes some of the cathode at high velocity and is accelerated out the nozzle by a Lorentz force. The thrust amount is controlled by varying the frequency of pulses with demonstrated range to date of 1–50 Hz producing thrust ranging from 1 µN to 0.05 mN. [ABSTRACT FROM AUTHOR]

We present current emission measurements from arrays of 480 porous ionic liquid ion source emitters for applications in space propulsion and in the processing and analysis of materials. The emitters are fabricated with a top-down approach from 10 × 10 × 1 mm3 bulk porous nickel and are housed within a microetched silicon frame and stainless-steel aperture-matching extraction grids. Measurements of emitted and collected currents are reported using two ionic liquids (ILs), EMI-BF4 and EMI-Im. Total beam currents up to about 400 µA can be sustained through passive feeding with IL fed from the rear of the bulk porous substrate. These currents are in agreement with previous work where currents of about 1 µA were observed from single emitters. Ion transparencies as high as 0.95 are achieved with manually assembled grids. Current emission displays reasonable symmetry between polarities, although slight differences could yield to the accumulation of one type of ion in the source leading to electrochemical reactions despite voltage alternation. Strong current decays with time and evidence of surface contamination reinforce this hypothesis. Future work should address this issue through closed-loop control of charge emission. [ABSTRACT FROM AUTHOR]

A conical pyrex plasma source with a conical helicon antenna surrounded by two cylindrically wound solenoids has been used in a helicon double layer thruster (HDLT) arrangement with the aim of investigating a new plasma cavity geometry and imposing a misalignment between the geometric and magnetic axis. An ion beam is measured with an energy analyser placed 6 cm downstream of the source exit for a low pressure argon plasma ([?]0.4 mTorr) for various currents in the two solenoids. Operation with current in the exhaust solenoid only also shows the presence of the ion beam. For a given radiofrequency power and similar total solenoidal current, the ion beam current measured downstream of the conical HDLT is about seven times greater than that previously measured downstream of the cylindrical HDLT. This may result from a larger plasma density in the source and from a source peak density shift by about 10 cm towards the tube exit measured in the conical arrangement compared with previous observations in the cylindrical arrangement. When an angle between the geometric and magnetic axis of up to 9deg is mechanically imposed, the ion beam is detected for angles less than 5deg. [ABSTRACT FROM AUTHOR]

We investigate the use of a centre-triggered cathodic arc as a spacecraft propulsion system that uses an inert solid as a source of plasma. The cathodic vacuum arc produces almost fully ionized plasma with a high exhaust velocity (>104 m s[?]1), giving a specific impulse competitive with other plasma or ion thrusters. A centre trigger design is employed that enables efficient use of cathode material and a high pulse-to-pulse repeatability. We compare three anode geometries, two pulse current profiles and two pulse durations for their effects on impulse generation, energy and cathode material usage efficiency. Impulse measurement is achieved through the use of a free-swinging pendulum target constructed from a polymer material. Measurements show that impulse is accurately controlled by varying cathode current. The cylindrical anode gave the highest energy efficiency. Cathode usage is optimized by choosing a sawtooth current profile. There is no requirement for an exhaust charge neutralization system. [ABSTRACT FROM AUTHOR]