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.

Time-resolved ion energy measurements using a retarding potential analyzer for electric propulsion applications.

To completely characterize the evolving state of a plasma, diagnostic tools that enable measurements of the time-resolved behavior are required. In this study, a gridded ion source with superimposed oscillations was utilized to verify the functionality of a high-speed retarding potential analyzer (HSRPA), at frequencies equivalent to the low frequency oscillations occurring in Hall effect thrusters (HETs). The verification of this device provides an effective alternative to existing diagnostics for measuring time-resolved ion energies. Retarding potential analyzers (RPAs) have established themselves as a fundamental diagnostic in the field of electric propulsion (EP), enabling the measurement of ion energy distributions within the plumes of EP thrusters. The work presented here has demonstrated the capability of a standard RPA in conjunction with high-speed circuitry and data fusion techniques to produce time-resolved ion energy distribution functions (IEDFs) at higher frequencies and beam potentials than have previously been investigated. Tested frequencies ranged between 20 and 80kHz with 10V peak-to peak oscillations at a mean beam potential of 570V. In addition, measurements were conducted with several waveforms, functioning as the superimposed oscillation, including a sine wave, triangle wave, and noisy sine wave. Data from the HSRPA were successfully reconstructed into time series utilizing two data fusion techniques: the empirical transfer function method and shadow manifold interpolation. Time-resolved IEDFs were produced at all frequency set points and waveforms. This investigation has demonstrated the HSRPA effectiveness at producing time-resolved measurements under conditions similar to those occurring in HETs.

Aromatic hydrocarbons as Molecular Propellants for Electric Propulsion Thrusters

The aromatic hydrocarbons (AHs) fluorobenzene, naphthalene, and 1-fluoronaphthalene are introduced as promising alternatives to xenon as propellant for in-space electric propulsion (EP). These storable molecules have similar mass, lower cost, and lower ionization energies compared to xenon, as well as the critical advantage of low post-ionization fragmentation compared to other molecular propellant candidates. The ionization characteristics of AHs are compared with those of xenon and the diamondoid adamantane, previously evaluated as a molecular propellant for EP. Quantum chemical calculations and BEB theory together with 25 eV electron ionization mass spectrometry (EI-MS) measurements have been used to predict the fragmentation of the AHs and adamantane when ionized in a plasma with an electron temperature of 7 eV (a typical electron temperature in EP plasmas). A high fraction (81 − 86%) of the detected AH ions originate from intact molecules, compared to 34% for adamantane, indicating extraordinarily low fragmentation for the selected AHs. The ionization potential of the AHs is similar to that of adamantane but lower compared to xenon (8.14–9.2 eV for the AHs, 9.25 for adamantane and 12.13 eV for xenon). BEB calculations have also been used to predict total ionization cross sections. The calculated ionization cross section of the AHs is comparable to that of adamantane but 3–5 times higher than that of xenon, which together with the low ionization potential can contribute to more efficient ionization. The AHs may have the potential to perform better than xenon, despite the absence of fragmentation in xenon.

Multiple fault isolation method for micro thrusters of drag-free systems

This paper proposes a multiple fault isolation framework for micro thrusters of drag-free systems, based on the minimum number of isolation filters. Based on the consideration of fault isolatability, the original system is reasonably divided into two completely isolated subsystems, and then two sets of fault isolation filters are designed based on the principles of duality design for state feedback decoupling, which achieves residual decoupling, and therefore, can detect and isolate multiple faults occurred on the micro thrusters. In particular, this method has low computational burden (a feature of practical importance in reducing the usage of onboard resources) and short isolation time, making it suitable for on-board fast and precise fault isolation. The simulation results demonstrate the efficacy of the proposed method when utilized for a drag-free system with 12 field emission electric propulsion (FEEP) thrusters.

Удосконалення методів нагріву ксенону для запобігання потраплянню рідкої фази робочої речовини до системи подачі

Among the various types of electric propulsion, the Hall thruster type is becoming the most common. This is due to the fact that the use of a Hall thruster makes it possible to obtain high values of the thruster characteristics with a simple design compared to other types of space propulsion systems. For Hall electric propulsion thrusters, the main working substance is xenon because of its fairly high atomic weight, low ionization energy, and unreactiveness, which makes it possible to obtain high thruster characteristics with ease of operation. The use of xenon as a working substance features a peculiarity involving its critical temperature (289.74 K), which gives rise to the liquid phase in the tank and, accordingly, pressure jumps, thus making it impossible to use the xenon feed system. To exclude the ingress of the liquid phase of xenon into the accumulator tank in electric propulsion systems, heaters are placed on the xenon tank to maintain its temperature within a given range. However, this approach has the following disadvantages: the low thermal conductivity of composite tanks impairs heater-to-xenon heat transfer; warming up the whole of the tank before starting the thruster increases the thruster start-up preparation time; the continuous maintenance of the tank temperature increases energy consumption by the propulsion system; and it is impractical to maintain the temperature of the whole of the xenon, while only a few grams of it are consumed for one thruster start-up. The problem that was solved in this work consists in changing the approach to heating the working substance that enters the feed system. The analysis of literary sources showed that this problem is relevant and offers ways to improve existing methods. To solve this problem, theoretical calculations were carried out and verified by experiment. As a result, a method was proposed to calculate the gasifier so that it may maintain the temperature of the working substance entering the accumulator tank within the range from 293 K to 298 K, thus eliminating the possible ingress of the liquid phase of xenon into the accumulator tank of the feed system. This study allows one to use the proposed structural element (gasifier) instead of tank heaters, which significantly reduces power consumption and maintains the stable operation of the working substance feed system. The conclusions drawn from the study may be useful to most developers of storage and feed systems for electric propulsion systems.

Progress in electric propulsion numerical simulation

Electric propulsion has been developed since the early 1960s, and its use onboard satellites, orbiting platforms, and interplanetary probes have increased significantly in the 21st century. The need for a detailed understanding of the working physics and a more accurate assessment of performance to create innovative designs has stimulated the development of several numerical simulation codes. The choice of method for modelling a specific thruster should be dictated by the physical characteristics of the flow in the device, and by the level of accuracy required from the simulation. There are various conditions in different types of thrusters. This means that different methods and computer codes must be developed for each of the different thrusters. The successful development of physically accurate numerical methods for simulating gas and plasma flows in electric propulsion thrusters can significantly improve the design process of these devices. In recent years, numerical simulations have increasingly benefited the basic understanding and engineering optimization of electric thrusters. This is due to several concurrent contributions: the evolution of computer hardware that has allowed the representation of multidimensional geometries and multiscale phenomena; implementation of sophisticated new algorithms and numerical diagnostic tools; and availability of new collisional and surface interaction data. There are two main directions for future work to continue to improve the numerical modelling of electric thrusters. First, the numerical methods themselves must be improved in terms of their physical accuracy and computational speed. The second main direction for improvement in the simulations involves more accurate determination of physical parameters that are required by the numerical formulations. This paper outlines efforts to develop models of various electrical propulsion concepts, from the first attempts in the early 90s to the latest sophisticated multidimensional simulations.

Thrust Pointing Adjustment of Electric Propulsion System

The satellite uses an electric propulsion system (EPS) to complete the north-south station-keeping (NSSK) work. Due to the change of satellite centroid position during the whole life period, electric propulsion thrust pointing adjustment is required to minimize the deviation between the thrust vector and the satellite centroid, to avoid the additional interference torque and the decrease of the position protection efficiency of the EPS. In this paper, the electric propulsion thrust pointing adjustment method is designed, the mathematical model of the pointing adjustment angle is established, the conformity of the thrust pointing adjustment during the whole life of the spacecraft is analyzed, and the ground integration test verification and the on-orbit electric propulsion thrust pointing adjustment verification are completed. The analysis and verification data show that the adjustment range and accuracy of the EPS can meet the requirements of the NSSK mission.

Semi-anechoic chamber for electromagnetic compatibility tests of electric propulsion thrusters

The present work deals with electromagnetic compatibility (EMC) investigations in the field of electric propulsion (EP). A semi-anechoic chamber (SAC) was set up to perform EMC tests with thrusters under operating conditions. The facility includes a semi-anechoic room and a vacuum tank. The measurement electronics currently have an upper frequency limit at 18 GHz. Initial measurements were done operating a Pulsed Plasma Thruster (PPT) to familiarize with the electronic systems. The same device under test (DUT) was investigated in a reverberation chamber (RVC) to obtain comparative results which showed good agreement. Furthermore, the anechoic chamber shows satisfactory initial measurement characteristics, which will be further tested and improved in the future. Characteristics such as chamber pressure, base noise show realistic values. The next steps are to perform measurements on radio-frequency ion thrusters (RIT) including peripherals and cabling.

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.

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.

Sub-millinewton thrust stand and wireless power coupler for microwave-powered small satellite thrusters.

The design and performance of a thrust stand for characterizing low-power electric propulsion thrusters are presented. The thrust stand is capable of sub-millinewton resolution for devices on the order of 1kg. The architecture is based on a counter-weighted hanging pendulum design, a variant of the standard hanging pendulum that employs a counterweight to increase force resolution. Thrust is measured in a displacement mode using the change in position of the pendulum arm as measured by an optical displacement sensor. Passive eddy-current damping is used to offset oscillations and decrease setting time. An in situ calibration rig using known masses is used to calculate thrust. The thrust stand features an adjustable counterweight for in-vacuum sensitivity adjustment. In addition, the design of a broadband (600-2490MHz) wireless microwave power coupler is presented. The device eliminates stiffness and thermal drift introduced by coaxial cables-typically the leading source of error in testing low-power microwave and radio frequency-powered thrusters. The thrust stand and coupler were tested using an electron cyclotron resonance magnetic nozzle thruster operating with xenon at flow rates from 1 to 10 sccm and powers ranging from zero (cold gas thrust) to 40W. The resulting measurements showed a force resolution of ∼10μN over a range of thrusts from ∼14 to 600µN.

Characteristics of plasma in a novel laser-assisted pulsed plasma thruster

A novel laser-assisted pulsed plasma thruster (LA-PPT) is proposed as an electric propulsion thruster, which separates laser ablation and electromagnetic acceleration. It aims for a higher specific impulse than that achieved with conventional LA-PPTs. Owing to the short-time discharge and the novel configuration, the physical mechanism of the discharge is unclear. Time and spatial-resolved optical emission spectroscopy was applied to investigate the variation in the plasma properties in the thruster discharge channel. The plasma species, electron temperature, and electron density were obtained and discussed. Our investigation revealed that there were H α , H β , H γ , H ε atoms, C I, C II, C III, C IV, Cl I, Cl II particles, and a small amount of CH, C3, C2, H2 neutral molecular groups in the plasma. The electron temperature of the discharge channel of the thruster was within 0.6–4.9 eV, and the electron density was within (1.1–3.0) 1018 cm−3, which shows that the optical emission spectroscopy method is to measure the electron excitation temperature and electron density in heavy particles. But the Langmuir probe method is to measure the temperature and density of free electrons. The use of laser instead of spark plug as the ignition mode significantly changed the plasma distribution in the discharge channel. Unlike the conventional PPT, which has high electron density near the thruster surface, LA-PPT showed relatively large electron density at the thruster outlet, which increased the thruster specific impulse. In addition, the change in the ignition mode enabled the electron density in the LA-PPT discharge channel to be higher than that in the conventional PPT. This proves that the ignition mode with laser replacing the spark plug effectively optimised the PPT performance.

Active attitude control for microspacecraft; A survey and new embedded designs

Novel technologies and design techniques of the active attitude control systems (ACS) for microspacecraft are briefly covered in this survey. Electric propulsion (EP) thrusters for satellites orbit correction and interplanetary vehicles acceleration have recently emerged as the focal point and front subject in propulsion fields due to their low mass, low cost and high specific impulse. Therefore, micro propulsion thrusters are explored for small satellites in terms of its performance, lifetime and future prospects. The focus of the review is at the center of research published in the last few years and state-of-the-art design concepts in terms of control, reliability, fault analysis and response to failure has been discussed for reaction wheels and magnetorquers. Finally, novel embedded planar magnetorquers which are integrated inside the internal layers of PCB for providing attitude maneuverability to various small satellites are discussed and simulated for critical tasks such as damping the initial angular velocities after vehicle launch. As the proposed magnetorquers are integrated in the form of copper traces, thermal analyses are done to maintain the temperature gradients.

In situ erosion measurement tools for electric propulsion thrusters: triangular laser head and telemicroscope

The lifetime of electric propulsion (EP) thrusters depends particularly on the erosion characteristics of operation relevant components, for instance, the grid hole erosion of gridded ion thrusters or the channel wall erosion of Hall effect thrusters. Here two tools for in situ erosion measurements are presented, a triangular laser head for surface profiling and a telemicroscope for high-resolution optical imaging. Both can give access to radial and axial erosion parameters. The measurements can be done in situ without the need for breaking the vacuum and dismounting the thruster, which reduces thruster testing time considerably. In situ measurements can also help to ensure reproducibility of thruster performance conditions and can improve statistics of thruster characterization. The present work describes the fundamentals of both techniques in detail, selected experimental setups are presented, their performance is characterized and critically evaluated. The capabilities and limitations related to erosion measurements of EP thrusters are, exemplary, demonstrated for a gridded ion thruster RIT-22 and a Hall effect thruster SPT-100D.

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.

TEST RESULTS OF A HIGH-SPEED SOLENOID VALVE FOR THE ELECTRIC PROPULSION FEED SYSTEM

Today in the world there are many electric propulsion systems based on the different types of thrusters. Most of them used gaseous propellent. Typical Electric propulsion system (EPS) consists of following main subsystems: the electric propulsion thruster that creates thrust; the power-processing and control unit which provides energy to all subsystems; the storage and feed system that performs functions of storing working substance in the tank and supplying the required amount of working substance to the thruster; Solenoid valve is a part of storage and feed systems. The feed system is an essential part of any electric propulsion system. The life cycle of the feed system is largely determined by lifetime of the valves and the amount of working substance in tank. Whereas the amount of working substance is selected based on the requirements for each specific mission valve parameters must be satisfying for various missions. Accordingly, the valves must have low power consumption, high reliability with a long lifetime (more than 106 cycles), low weight and dimensions. Solenoid valve is an electromechanical device that consists of moving part controlled by an electromagnet. Movement of the movable element and the seal are key parts of valve reliability. Presented solenoid valve design uses two membranes, which perform the function of centering the moving element and with the spring provide the required closing speed and the seal of the valve. The paper presents the problem of selecting critical elements that affect the solenoid valve's performance, procedure and test results required for the qualification of space equipment. Manufactured valve was successfully tested (vibration testing, response time testing with inlet pressure change, response time testing at different operating voltages, dependence of the valve actuation current on the operating voltage, temperature testing, lifetime tests) separately and as part of XFS according to the ECSS standards. As well as the successful implementation of presented valve into the flight xenon feed system for electric propulsion.

Control of a service satellite during its mission on space debris removal from orbits with high inclination by implementation of an ion beam method

The largest population of space debris is observed in the low-Earth-orbit region with altitudes of 600–1000 km and inclinations of 80–100°. As follows from the predictive calculations, space debris removal from this region exactly is of the first priority. For this reason, our article discusses the problems of controlling the motion of service spacecraft – space debris object cluster when the space debris object is removed from a low orbit with high inclination. In the conceptual design of the spacecraft presented hereinafter, the electric propulsion system contains two electric propulsion thrusters of stationary plasma thruster type, each of which is mounted with the possibility of independent rotation respect to two mutually perpendicular directions. An ion beam injector forms an ion beam with 65 mN of thrust. Spacecraft motion control, including the thrust vector control, is provided by coordinated rotation of thrusters. For such concept, a strategy and an algorithm for controlling the motion of the spacecraft center of mass are proposed, which are based on the control of the spacecraft lateral and longitudinal motion that is understood as the control of the relative distance between the spacecraft and the space debris object. The strategy of spacecraft lateral motion control assumes that due to the thrusters' rotation the spacecraft should be displaced in such a way that the vector of the relative range between the spacecraft and the space debris object coincides with the direction of the transversal of the orbital coordinate system. For the practical implementation of the control, an algorithm has been developed for controlling the thrusters' rotation to create the required values of the spacecraft acceleration projections. The mathematical model includes the equations of motion of the spacecraft center of mass taking into account the aerodynamic deceleration. The solar panels' orientation to the Sun is provided by the spacecraft roll rotation relative to its longitudinal axis and turning the solar panels relative to their axis of rotation. The change in the ion beam momentum transfer is considered using a simple simulation model. Mathematical modeling was performed for the following parameters of the initial orbit: altitude: 800 km; eccentricity: 0.0005; and inclination: 900. For the proposed control algorithm, a range of possible thrust of a single thruster was 51–60 mN. The simulation results show that the proposed control can be used to remove space debris object from low orbits with high inclination, taking into account the roll spacecraft rotation to keep the solar panels’ orientation to the Sun.

КВАЗІОДНОВИМІРНА МАТЕМАТИЧНА МОДЕЛЬ ПРОЦЕСІВ В ХОЛЛІВСЬКОМУ ТА ПЛАЗМОВО-ІОННОМУ ДВИГУНІ

Plasma-ion thrusters with a radial magnetic field in ionization chamber and Hall effect thrusters are electrostatic electric propulsion thrusters with closed electron drift. Axial symmetry in the dynamics of the components of propellant in these thrusters allows to write the equations of plasma-dynamics for electrons, ions and neutral atoms in two-dimensional axial-radial form. Attempts to reduce the equations to simpler one-dimensional form by simply removing the components with radius differentiation lead to the loss in the description of important effects, responsible for values of thruster performance. At the same time, a significant disadvantage of gas dynamics equation set is its fundamental openness – the correspondence between the number of unknown variables and equations is achieved approximately basing on some assumptions. In traditional form of gas dynamics, such closeness is made under the assumption of thermodynamic equilibrium with velocity distribution functions of components close to Maxwell one, which is the limit result of collisions. The use of such approximation to plasma components dynamics in electric propulsion thrusters is impossible due to the rarefaction of the substance in them. A mathematical model of two-component plasma-dynamics is represented in stationary form to describe the processes in the Hall effect thruster channel and the ionization chamber of plasma-ion thruster with radial magnetic field. Due to the impossibility of using the method of local thermodynamic equilibrium to describe the rarefied substance in electric propulsion thruster, a more advanced form of equations is used. A more reliable means of approximate closure of the set of equations is proposed in the description of the rarified gas. An approach to the description of the specifics of electrons energy transfer from the plasma to the walls of the channel, as well as the non-mirror reflection of electrons from the potential barrier within the Langmuir layer is shown. A method of averaging the parameters over the cross section of the channel is proposed, which allows to convert the equation into a quasi-one-dimensional form with the preservation of charge, momentum and energy losses on the channel walls.

The ablation characteristics of laser-assisted pulsed plasma thruster with metal propellant

In this study, a laser-assisted pulsed plasma thruster (LA-PPT) with a novel configuration is proposed as an electric propulsion thruster which separates laser ablation and electromagnetic acceleration. Owing to the unique structure of the thruster, metals can also be used as propellants, and a higher specific impulse is expected. The ablation quality, morphology, and plume distribution of various metals (aluminium alloy, red copper, and carbon steel) with different laser energies were studied experimentally. The ablation morphology and plume distribution of red copper were more uniform, as compared to those of other metals, and the ablation quality was higher, indicating its greater suitability for LA-PPT. The plume generated by nanosecond laser ablation of aluminium alloy expanded faster, which indicated that the response time of the thruster with aluminium alloy as the propellant was shorter. In addition, when the background pressure was 0.005 Pa, an obvious plume splitting phenomenon was observed in the ablation plume of the pulsed laser irradiating aluminium alloy, which may significantly reduce the utilisation rate of the propellant.

Radial distribution of electrons rotation moment in hall effect and plasma-ion thrusters

The subject matter of the article is the radial distribution of electrons movement parameters inside electric propulsion thrusters with closed electrons drift. The radial magnetic field in Hall effect thrusters is the limits the axial flow of electrons because of interaction with azimuth electron current. In turn, this azimuth current exists as a result of rivalry between the attempt of the magnetic field to transform electrons current completely closed one and the loss of electrons rotation moment in collisions. Similar processes take place in the ionization chamber of plasma-ion thrusters with the radial magnetic field. The attempts to estimate electrons parameters through only collisions with ions and atoms inside volume have given the value of axial electrons current much lower than really being. This phenomenon is called anomalous electrons conductivity, which was tried to be explained as a consequence of various effects including "near-the-wall-conductivity", which was explained as a result of non-mirror reflection of electrons from the Langmuir layer near the walls of the thruster channel. The disadvantage of this name is the fact that the reflection of the electron occurs before reaching the surface from the potential barrier at the plasma boundary with any environment: the wall, but also with the environment vacuum. The potential distribution in the Langmuir layer is non-stationary and inhomogeneous due to the presence of so-called plasma oscillations. The definition of "conductivity" is just as unfortunate in this name, because the collisions are always not a factor of conductivity, but on the contrary – of resistance. The goal is to solve the task of electrons rotation moment distribution in the thruster channel. The methods used are the formulation of the kinetic equation for electrons distribution function over the velocities, radius, and projections of the coordinates of the instantaneous center of cyclotron rotation; solution of this equation and finding with its use the distribution of the gas-dynamic parameters of electrons along the cross-section of the channel. Conclusions. A mathematical model of electrons rotation moment dynamics is proposed, which allows using plasma-dynamics equations to analyze its distribution along the cross-section of thruster channel and to estimate the effect of "near-the-wall-conductivity" using appropriate boundary conditions.