Electric jet engine. What is active and reactive electricity? See what "Electric rocket engine" is in other dictionaries

Plan:

1. Introduction
2 Classification of electric propulsion
3 Brief technical specifications
4 History
5 Perspectives

Introduction

Electric rocket engine (ERE)- a rocket engine, the operating principle of which is based on the conversion of electrical energy into kinetic energy of particles. There are also names that include the words reactive And mover.

A complex consisting of a set of electric propulsion engines, a working fluid storage and supply system (SHiP), an automatic control system (ACS), and a power supply system (SPS) is called electric propulsion system (EPS).

1. Introduction

The idea of using electrical energy to accelerate the working fluid (PM) in jet engines arose almost at the beginning of the development of rocket technology. It is known that such an idea was expressed by K. E. Tsiolkovsky. In 1916-1917, R. Goddard conducted the first experiments, and in the 30s of the 20th century in the USSR, under the leadership of V.P. Glushko, one of the first operating electric propulsion engines was created.

From the very beginning, it was assumed that the separation of the energy source and the accelerated substance would ensure a high speed of exhaustion of the RT, as well as a lower mass of the spacecraft (SC) due to a decrease in the mass of the stored working fluid. Indeed, in comparison with other rocket engines, electric propulsion engines make it possible to significantly increase the active lifetime (AS) of a spacecraft, while significantly reducing the mass of the propulsion system (PS), which, accordingly, makes it possible to increase the payload or improve the weight-dimensional characteristics of the spacecraft itself.

Calculations show that the use of electric propulsion will reduce the duration of flights to distant planets (in some cases even make such flights possible) or, with the same flight duration, increase the payload.

Starting from the mid-60s, full-scale tests of electric propulsion engines began in the USSR and the USA, and in the early 70s, electric propulsion engines began to be used as standard propulsion systems.

Currently, electric propulsion systems are widely used both in the propulsion systems of Earth satellites and in the propulsion systems of interplanetary spacecraft.

2. Classification of electric propulsion engines

The classification of electric propulsion engines has not been established, however, in Russian-language literature it is usually customary to classify electric propulsion engines according to predominant particle acceleration mechanism. The following types of engines are distinguished:

electrothermal rocket engines (ETR);
electrostatic motors (ID, SPD);
high-current (electromagnetic, magnetodynamic) motors;
impulse motors.

Classification of electric rocket engines accepted in Russian-language literature

ETDs, in turn, are divided into electric heating (END) and electric arc (EDA) engines.

Electrostatic engines are divided into ion (including colloidal) engines (ID, CD) - particle accelerators in a unipolar beam, and particle accelerators in a quasineutral plasma. The latter include accelerators with closed electron drift and an extended (UZDP) or shortened (UZDU) acceleration zone. The first are usually called stationary plasma engines (SPD), and the name is also found (increasingly less often) - linear Hall engine (LHD), in Western literature it is called a Hall engine. Ultrasonic motors are usually called anode-accelerated motors (LAMs).

High-current (magnetoplasma, magnetodynamic) motors include motors with their own magnetic field and motors with an external magnetic field (for example, an end-mounted Hall motor - THD).

Pulse engines use the kinetic energy of gases produced by the evaporation of a solid in an electrical discharge.

Any liquids and gases, as well as their mixtures, can be used as a working fluid in electric propulsion engines. However, for each type of engine there are working fluids, the use of which allows you to achieve the best results. For ETD, ammonia is traditionally used, for electrostatic - xenon, for high-current - lithium, for pulse - fluoroplastic.

The disadvantage of xenon is its cost, due to its small annual production (less than 10 tons per year worldwide), which forces researchers to look for other RTs with similar characteristics, but less expensive. Argon is considered as the main candidate for replacement. It is also an inert gas, but, unlike xenon, it has higher ionization energy with a lower atomic mass. The energy spent on ionization per unit of accelerated mass is one of the sources of efficiency losses.

3. Brief technical specifications

Electric propulsion engines are characterized by a low RT mass flow rate and a high outflow velocity of an accelerated particle flow. The lower limit of the exhaust velocity approximately coincides with the upper limit of the exhaust velocity of a chemical engine jet and is about 3,000 m/s. The upper limit is theoretically unlimited (within the speed of light), however, for promising engine models, a speed not exceeding 200,000 m/s is considered. Currently, for engines of various types, the optimal exhaust velocity is considered to be from 16,000 to 60,000 m/s.

Due to the fact that the acceleration process in an electric propulsion engine takes place at low pressure in the accelerating channel (particle concentration does not exceed 10 20 particles/m³), the thrust density is quite low, which limits the use of electric propulsion engines: the external pressure should not exceed the pressure in the accelerating channel, and the acceleration of the spacecraft is very small (tenths or even hundredths g ). An exception to this rule may be EDD on small spacecraft.

The electrical power of electric propulsion engines ranges from hundreds of watts to megawatts. Electric propulsion engines currently used on spacecraft have a power from 800 to 2,000 W.

Electric propulsion engines are characterized by not very high efficiency - from 30 to 60%.

4. History

In 1964, in the attitude control system of the Soviet Zond-2 spacecraft, 6 erosive pulse thrusters operating on fluoroplastic operated for 70 minutes; the resulting plasma clots had a temperature of ~ 30,000 K and flowed out at a speed of up to 16 km/s (the capacitor bank had a capacity of 100 μF, the operating voltage was ~ 1 kV). In the USA, similar tests were carried out in 1968 on the LES-6 spacecraft. In 1961, a pinch pulse taxiway from the American company Republic Aviation (eng. Republic Aviation) developed a thrust of 45 mN on the stand with a specific impulse of 10-70 km/s. In 1971, two end Hall thrusters operated in the correction system of the Soviet Meteor satellite, each of which, with a power supply of ~ 0.5 kW, developed a thrust of 18-23 mN and a specific impulse of over 8 km/s. The RDs had a size of 108×114×190 mm, a mass of 32.5 kg and a reserve of Xenon (compressed xenon) of 2.4 kg. During one of the switch-ons, they worked continuously for 140 hours.

5. Prospects

Currently, many countries are exploring the creation of manned interplanetary spacecraft with electric propulsion systems. Existing electric propulsion engines are not optimal for use as propulsion engines for such ships, and therefore in the near future we should expect renewed interest in the development of high-current electric propulsion engines based on liquid metal RT (bismuth, lithium, potassium, cesium) with an electrical power of up to 1 MW, capable work for a long time at currents up to 5-10 kA. These taxiways must develop a thrust of up to 20-30 N and a specific impulse of 20-30 km/s with an efficiency of 30% or more. In 1975, a similar thruster was tested in the USSR on the Cosmos-728 satellite (a thruster with an electrical power of 3 kW, powered by potassium, developed a specific impulse of ~ 30 km/s).

In addition to Russia and the USA, research and development of electric propulsion systems is also carried out in the UK, Germany, France, Japan, and Italy. The main areas of activity of these countries: ID (the most successful developments are from Great Britain and Germany, especially joint ones); SPD and DAS (Japan, France); ETD (France). These engines are mainly intended for satellites.

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The idea of using electrical energy in jet engines for acceleration arose almost at the beginning of the development of rocket technology. It is known that such an idea was expressed by K. E. Tsiolkovsky. In -1917, R. Goddard conducted the first experiments, and in the 30s of the 20th century in the USSR, under the leadership of V.P. Glushko, one of the first operating electric propulsion engines was created.

From the very beginning, it was assumed that the separation of the energy source and the accelerated substance would provide a high speed of exhaust of the working fluid (PT), as well as a lower mass of the spacecraft (SC) due to a decrease in the mass of the stored working fluid. Indeed, in comparison with other rocket engines, electric propulsion engines make it possible to significantly increase the active lifetime (AS) of a spacecraft, while significantly reducing the mass of the propulsion system (PS), which, accordingly, makes it possible to increase the payload or improve the weight-dimensional characteristics of the spacecraft itself.

Classification of electric rocket engines accepted in Russian-language literature

ETDs, in turn, are divided into electric heating (END) and electric arc (EDA) engines.

Electrostatic engines are divided into ion (including colloidal) engines (ID, CD) - particle accelerators in a unipolar beam, and particle accelerators in a quasineutral plasma. The latter include accelerators with closed electron drift and an extended (UZDP) or shortened (UZDU) acceleration zone. The first ones are usually called stationary plasma engines (SPD), and the name also appears (increasingly less often) - linear Hall engine (LHD), in Western literature it is called a Hall engine. Ultrasonic motors are usually called anode-accelerated motors (LAMs).

These include motors with their own magnetic field and motors with an external magnetic field (for example, an end-mounted Hall motor - THD).

Pulse engines use the kinetic energy of gases produced by the evaporation of a solid in an electrical discharge.

Any liquids and gases, as well as their mixtures, can be used as a working fluid in electric propulsion engines. However, for each type of engine there are working fluids, the use of which allows you to achieve the best results. Ammonia is traditionally used for ETD, xenon for electrostatic, lithium for high-current, and fluoroplastic for pulsed.

The disadvantage of xenon is its cost, due to its small annual production (less than 10 tons per year worldwide), which forces researchers to look for other RTs with similar characteristics, but less expensive. Argon is being considered as the main candidate for replacement. It is also an inert gas, but, unlike xenon, it has higher ionization energy with a lower atomic mass. The energy spent on ionization per unit of accelerated mass is one of the sources of efficiency losses.

The electrical power of electric propulsion engines ranges from hundreds of watts to megawatts. Electric propulsion engines currently used on spacecraft have a power from 800 to 2,000 W.

Electric jet engine in the Polytechnic Museum, Moscow. Created in 1971 at the Institute of Atomic Energy named after. I. V. Kurchatova

On October 1, 1966, the Yantar-1 automatic ionospheric laboratory was launched to an altitude of 400 km by a three-stage geophysical rocket 1YA2TA to study the interaction of the jet stream of an electric rocket engine (ERE), running on argon, with ionospheric plasma. The experimental plasma-ion electric propulsion engine was first turned on at an altitude of 160 km, and during the subsequent flight 11 cycles of its operation were carried out. A jet stream velocity of about 40 km/s was achieved. The Yantar laboratory reached a specified flight altitude of 400 km, the flight lasted 10 minutes, the electric propulsion engine operated steadily and developed a design thrust of five grams of force. The scientific community learned about the achievement of Soviet science from a TASS report.

In the second series of experiments, nitrogen was used. The exhaust speed was increased to 120 km/s. In 1971, four similar devices were launched (according to other sources, before 1970 there were six devices).

In the fall of 1970, a ramjet electric propulsion system successfully passed tests in real flight. In October 1970, at the XXI Congress of the International Astronomical Federation, Soviet scientists - Professor G. Grodzovsky, Candidates of Technical Sciences Yu. Danilov and N. Kravtsov, Candidates of Physical and Mathematical Sciences M. Marov and V. Nikitin, Doctor of Technical Sciences V. Utkin - reported on testing of an air propulsion system. The recorded jet speed reached 140 km/s.

In 1971, the correction system of the Soviet meteorological satellite “Meteor” operated two stationary plasma engines developed by the Fakel Design Bureau, each of which, with a power supply of ~ 0.4 kW, developed a thrust of 18-23 mN and an exhaust velocity of over 8 km/s. The RDs had a size of 108×114×190 mm, a mass of 32.5 kg and a reserve of Xenon (compressed xenon) of 2.4 kg. During one of the starts, one of the engines worked continuously for 140 hours. This electric propulsion system is shown in the figure.

Electric rocket engines are also used in the Dawn mission. Planned use in the BepiColombo project.

Although electric rocket engines have low thrust compared to liquid-fuel rockets, they are capable of operating for long periods of time and capable of slow flight over long distances.

The invention relates to electric jet engines. The invention is an end-type engine on a solid working fluid, consisting of an anode, a cathode and a working fluid block located between them. The block is made of a material with a high dielectric constant, such as barium titanate, and an anode and a cathode are installed on one side, and a conductor is attached to the other side. The checker can be in the shape of a disk with a cathode and anode installed coaxially or diametrically opposite. The invention makes it possible to create a pulsed electric jet engine of simple design with high specific parameters. 4 salary f-ly, 2 ill.

The invention relates to the field of electric jet engines (EPM) of pulse action on a solid-phase working fluid. Pulse plasma engines with a gaseous working fluid supply system (for example, xenon, argon, hydrogen) and erosion-type pulse engines with a solid-phase working fluid polytetrafluoroethylene (PTFE) are known. The main disadvantage of the first type of engine is the complex system of pulsed, strictly dosed supply of the working fluid due to the difficulty of synchronizing it with discharge voltage pulses and, as a consequence, the low utilization rate of the working fluid. In the second case (erosive type, working fluid - PTFE), the specific parameters have low values, the maximum efficiency does not exceed 15% due to the prevailing thermal mechanism of producing and accelerating the electric discharge plasma. A more advanced type of engine of this class is an end-type pulsed electric plasma jet engine on a solid working fluid (including PTFE) with a predominant electron-detonation type of breakdown (explosive injection of electrons from the surface of the working fluid towards the anode). This type of engine makes it possible to obtain higher specific parameters using the PTFE working fluid due to a significant reduction in the arc phase of the plasma source discharge. The presence of the arc stage of the discharge also leads to the appearance of instability in the plasma generation process on the surface of the working fluid such as plasma bundles with the formation of channels with increased conductivity on the surface of the working fluid and, as a consequence, to short-circuiting the interelectrode gap along the mentioned channels. The literature describes the results of studies on the incomplete type of breakdown on the surface of a dielectric at currents realized at the moment of charging a capacitor containing a dielectric with a high dielectric constant. Based on this type of breakdown, an effective source of pulsed-type particles (ions or electrons) has been created. However, when assessing the possibility of using it as part of a pulsed electric propulsion engine based on an ion component with a switching frequency of tens to hundreds of hertz, problems arise with the discharge (depolarization) of the dielectric used as a working fluid, as well as problems with the durability of the grid electrode, which acts as a particle extractor, and problems neutralization of ions. The purpose of the proposed invention is to create a pulse electric propulsion engine that is simple in design with a switching frequency of up to 100 hertz or more to obtain low thrust per single discharge of the generator, but with high specific parameters. The desired level of traction second impulse is ensured by adjusting the switching frequency. This goal is achieved by the fact that in an end-type pulsed electric reluctance motor on a solid working fluid consisting of an anode, a cathode and a working fluid block located between them, it is proposed that the working fluid block be made of a dielectric with a high dielectric constant and installed on one side of the block anode and cathode, and install or apply a conductor on the other side of the checker. The preferred material for the working fluid block is barium titanate, and the most constructive form is the disk form. The anode and cathode can be installed coaxially or diametrically opposite. The proposed solution is illustrated by drawings. Figure 1 shows a variant of a pulsed electric propulsion engine with a coaxially located anode and cathode; Fig. 2 shows a variant with anode and cathode installed diametrically opposite. The proposed engine consists of an anode, a cathode and a working fluid block made of a dielectric with a high dielectric constant, for example barium titanate with 1000. Such a block can have the shape of a disk, on one side of which conductor 2 is applied in the form of a thin layer, for example, by spraying or in the form of a metal plate tightly pressed to the surface of the dielectric. On the other side of the checker there is an anode 3 and a cathode 4, located either coaxially (Fig. 1) or diametrically opposite (Fig. 2). In such a device, when voltage is applied to the anode and cathode, the interelectrode overlap of the dielectric occurs along the surface of the dielectric and begins from both electrodes as a result of charging two series-connected capacitors formed by the “anode - dielectric - conductor” and “conductor - dielectric - cathode” systems. As a result, we have two plasma torches (anode and cathode) above the surface of the dielectric, moving towards each other, while conductor 2 (conducting plate) of the device will have a floating potential, due to the nature of the flow of displacement currents through the dielectric. At the moment of merging of the anode and cathode torches, the excess positive charge of ions is neutralized, the formation mechanism of which is due to the electron-detonation type of breakdown for the anode torch. The plasma obtained after the fusion of two torches acquires additional acceleration in the mode of discharge (depolarization) and release of the energy stored in such a capacitor, similar to a linear accelerator. To realize the effect of additional acceleration, the height of the electrodes (anode and cathode) along the plasma flow is formed based on the real time required to discharge the capacitance of the electric propulsion engine design. This design of the device and its operating mode make it possible to create a pulsed electric propulsion engine with high parameter values and a high switching frequency (a prototype of the specified type of electric propulsion engine based on modified standard high-voltage (less than 10 kV) capacitors of the KVI-3 type operates at NIIMASH with a switching frequency of up to 50 Hz) . To operate such an electric propulsion engine, a generator of high-voltage pulses of nanosecond duration is required. The duration of the pulses supplied to the electrodes is determined by the charging time of the capacitance of the electric propulsion engine design. To eliminate instabilities such as plasma bundles, the duration of the high-voltage pulse from the generator should not exceed the duration of charging the capacitance of the electric propulsion engine design. The maximum switching frequency of the electric propulsion engine is determined by the time required for a full cycle of charging and discharging the capacity of the electric propulsion engine design. The dimensions of the cathode and anode plasma torches moving towards each other are determined by the dielectric overlap rate, which depends on the voltage amplitude, the value of the structure’s capacitance, as well as the delay time for the start of the plasma torch generation process. This delay time, in turn, depends on the geometric parameters of the anode-dielectric, cathode-dielectric zone, the type of dielectric, and the area of the conductor. This electric propulsion engine works as follows. When a high-voltage voltage pulse is applied to the anode 3 and cathode 4 with a duration corresponding to the charging time of the capacitance of the electric propulsion engine design, two plasma torches moving towards each other are generated (anode from the anode and cathode from the cathode). The anode torch has an excess positive charge of ions of the working fluid (in relation to such a dielectric as barium titanate ceramics, these are mainly barium ions as the most easily ionized element). The cathode plume plasma is caused by the generation of electrons from the cathode and their bombardment of the dielectric surface. At the moment of meeting, the cathode torch neutralizes the anode one and the plasma bunch is accelerated like a linear accelerator in the phase of discharging the capacity of the electric propulsion design through the plasma. It should be noted that the zones of inter-flame breakdowns that arise when flame torches approach each other are not strictly localized, that is, they are not “tied” to certain places on the surface of the dielectric during the production of a large number of pulses. The specified operating mode of such an electric propulsion engine will contribute to obtaining high efficiency values and plasma outflow rates. An essential feature of the proposed electric propulsion engine is the pulse-frequency operating mode (with a frequency of up to 100 Hz or more) with the ability to almost instantly gain and release thrust. Thanks to this feature and taking into account the electrical power actually available on board the spacecraft (SC), the area of effective application of the propulsion system (PS) based on the proposed pulsed electric propulsion system can be expanded, namely:

Maintaining geostationary spacecraft in the north-south, east-west direction;

Compensation of spacecraft aerodynamic drag;

Changing orbits and moving spent or failed spacecraft to a given area. Information sources

1. Grishin S.D., Leskov L.V., Kozlov N.P. Electric rocket engines. - M.: Mechanical Engineering, 1975, p. 198-223. 2. Favorsky O.N., Fishgoit V.V., Yantovsky E.I. Fundamentals of the theory of space electric propulsion systems. - M.: Mechanical Engineering, Higher School, 1978, p. 170-173. 3. L. Caveney (translation from English edited by A.S. Koroteev). Space engines - status and prospects. - M., 1988, p. 186-193. 4. Patent for invention 2146776 dated May 14, 1998. End-type pulsed plasma jet engine on a solid working fluid. 5. Vershinin Yu.N. Electron-thermal and detonation processes during electrical breakdown of solid dielectrics. Ural Branch of the Russian Academy of Sciences, Ekaterinburg, 2000. 6. Bugaev S.P., Mesyats G.A. Emission of electrons from the plasma of an incomplete discharge through a dielectric in a vacuum. DAN USSR, 1971, vol. 196, 2. 7. Mesyats G.A. Actons. Part 1-Ural Branch of the Russian Academy of Sciences, 1993, p. 68-73, part 3, p. 53-56. 8. Bugaev S.P., Kovalchuk B.M., Mesyats G.A. Plasma pulsed source of charged particles. Copyright certificate 248091.

CLAIM

1. An end-type pulsed electric reluctance motor on a solid working fluid, consisting of an anode, a cathode and a working fluid block made of a dielectric with a high dielectric constant and located between them, characterized in that the cathode and anode are located on one side of the block and are removed from each other, and a conductor is applied to the other side. 2. Pulse electric jet engine according to claim 1, characterized in that the working fluid block is made of barium titanate. 3. Pulse electric jet engine according to claim 1, characterized in that the working fluid block has the shape of a disk. 4. Pulse electric reluctance motor according to claim 3, characterized in that the cathode and anode are installed coaxially. 5. Pulse electric reluctance motor according to claim 3, characterized in that the cathode and anode are installed diametrically opposite.

Electric rocket motor

An electric rocket engine is a rocket engine whose operating principle is based on the use of electrical energy received from a power plant on board the spacecraft to create thrust. The main area of application is minor trajectory correction, as well as space orientation of spacecraft. A complex consisting of an electric rocket engine, a working fluid supply and storage system, an automatic control system and a power supply system is called an electric rocket propulsion system.

Mention of the possibility of using electric energy in rocket engines to create thrust is found in the works of K. E. Tsiolkovsky. In 1916-1917 The first experiments were carried out by R. Goddard, and already in the 30s. XX century under the leadership of V.P. Glushko, one of the first electric rocket engines was created.

In comparison with other rocket engines, electric ones make it possible to increase the lifespan of a spacecraft, and at the same time the weight of the propulsion system is significantly reduced, which makes it possible to increase the payload and obtain the most complete weight and size characteristics. Using electric rocket engines, it is possible to shorten the duration of flights to distant planets, and also make flights to any planet possible.

In the mid-60s. XX century Electric rocket engines were actively tested in the USSR and the USA, and already in the 1970s. they were used as standard propulsion systems.

In Russia, classification is based on the mechanism of particle acceleration. The following types of engines can be distinguished: electrothermal (electric heating, electric arc), electrostatic (ionic, including colloidal, stationary plasma engines with acceleration in the anode layer), high-current (electromagnetic, magnetodynamic) and pulse engines.

Any liquids and gases, as well as their mixtures, can be used as a working fluid. For each type of electric motor, it is necessary to use the appropriate working fluids to achieve the best results. Ammonia is traditionally used for electrothermal motors, xenon is used for electrostatic motors, lithium is used for high-current motors, and fluoroplastic is the most effective working fluid for pulse motors.

One of the main sources of losses is the energy spent on ionization per unit of accelerated mass. The advantage of electric rocket engines is the low mass flow of the working fluid, as well as the high speed of the accelerated flow of particles. The upper limit of the outflow velocity is theoretically within the speed of light.

Currently, for various types of engines, the exhaust velocity ranges from 16 to 60 km/s, although promising models will be able to give an exhaust velocity of the particle flow of up to 200 km/s.

The disadvantage is the very low thrust density; it should also be noted that the external pressure should not exceed the pressure in the acceleration channel. The electrical power of modern electric rocket engines used on spacecraft ranges from 800 to 2000 W, although theoretical power can reach megawatts. The efficiency of electric rocket engines is low and varies from 30 to 60%.

In the next decade, this type of engine will mainly perform tasks for correcting the orbit of spacecraft located in both geostationary and low-Earth orbits, as well as for delivering spacecraft from the reference low-Earth orbit to higher ones, such as geostationary orbit.

Replacing a liquid rocket engine, which serves as an orbit corrector, with an electric one will reduce the mass of a typical satellite by 15%, and if the period of its active stay in orbit is increased, then by 40%.

One of the most promising areas for the development of electric rocket engines is their improvement in the direction of increasing power to hundreds of megawatts and specific thrust impulse, and it is also necessary to achieve stable and reliable operation of the engine using cheaper substances, such as argon, lithium, nitrogen.

From the book Great Soviet Encyclopedia (AN) by the author TSB

From the book Great Soviet Encyclopedia (DV) by the author TSB

From the book Great Soviet Encyclopedia (RA) by the author TSB

From the book Great Soviet Encyclopedia (SB) by the author TSB

From the book Great Soviet Encyclopedia (SU) by the author TSB

From the book Great Soviet Encyclopedia (EL) by the author TSB

From the book Great Encyclopedia of Technology author Team of authors

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ELECTRIC ROCKET MOTOR, electric rocket engine(ERD) - rocket engine, in which the electrical energy of the spacecraft’s onboard power plant (usually solar or battery batteries) is used as an energy source to create thrust. According to the principle of operation, electric propulsion engines are divided into electrothermal rocket engines, electrostatic rocket motors And electromagnetic rocket engines. In electrothermal RDs, electrical energy is used to heat the working fluid (WM) in order to convert it into a gas with a temperature of 1000-5000 K; the gas flowing out of the jet nozzle (similar to the nozzle of a chemical rocket engine) creates thrust. In electrostatic jet engines, for example, ionic jets, the RT is first ionized, after which positive ions are accelerated in an electrostatic field (using a system of electrodes) and, flowing out of the nozzle, create thrust (to neutralize the charge of the jet stream, electrons are injected into it). In an electromagnetic RD (plasma), the working fluid is the plasma of any substance, accelerated due to the Ampere force in crossed electric and magnetic fields. Based on the indicated main types (classes) of electric propulsion engines, it is possible to create various intermediate and combined options that best meet the specific conditions of application. In addition, some electric propulsion engines can “transition” from one class to another when the power supply mode changes.

The electric propulsion engine has an exceptionally high specific impulse - up to 100 km/s or more. However, the large required energy consumption (1-100 kW/N of thrust) and the small ratio of thrust to the cross-sectional area of the jet stream (no more than 100 kN/m 2) limit the maximum expedient thrust of an electric propulsion engine to several tens of newtons. Electric propulsion engines are characterized by dimensions of ~0.1 m and a mass of the order of several kilograms.

The working fluids of electric propulsion engines are determined by the essence of the processes occurring in various types of these engines and are very diverse: these are low molecular weight or easily dissociating gases and liquids (in electrothermal thrusters); alkaline or heavy, easily evaporating metals, as well as organic liquids (in electrostatic RD); various gases and solids (in electromagnetic RD). Typically, the tank with the RT is structurally combined with the electric propulsion engine in a single propulsion unit (module). The separation of the energy source and the RT contributes to very precise control of the thrust of the electric propulsion engine over a wide range while maintaining a high specific impulse value. Many electric propulsion engines are capable of operating for hundreds and thousands of hours when switched on repeatedly. Some electric propulsion engines, which are pulsed propulsion engines by their principle, allow tens of millions of inclusions. The efficiency and perfection of the working process of electric propulsion are characterized by the values of the efficiency coefficient and traction prices, electric propulsion dimensions - value thrust density.

Characteristic values of some electric propulsion parameters

Options	Electric propulsion type
Options	electro-thermal	electromagnetic	electrostatic
Thrust, N	0,1 — 1	0,0001 — 1	0,001 — 0,1
Specific impulse, km/s	1 — 20	20 — 60	30 — 100
Thrust density (maximum), kN/m 2	100	1	0,03 — 0,05
Supply voltage, V	units - tens	tens - hundreds	tens of thousands
Supply current strength, A	hundreds - thousands	hundreds - thousands	fractions of a unit
Thrust price, kW/N	1 — 10	100	10 — 40
Efficiency	0,6 — 0,8	0,3 — 0,5	0,4 — 0,8
Electrical power, W	tens - thousands	units - thousands	tens - hundreds

An important characteristic of the electric propulsion engine is the power supply parameters. Due to the fact that most existing and future on-board power plants are characterized by the generation of direct current of relatively low voltage (units - tens of volts) and high power (up to hundreds and thousands of amperes), the easiest way to solve the issue of power supply is in electrothermal RDs, which are predominantly low-voltage and high current. These RDs can also be powered from an alternating current source. The greatest difficulties with power supply arise when using electrostatic RDs, the operation of which requires a direct current of high (up to 30-50 kV) voltage, although of low strength. In this case, it is necessary to provide conversion devices that significantly increase the mass of the remote control. The presence in the propulsion system of working elements associated with the electric propulsion power supply and the low value of the electric propulsion thrust determine the extremely low thrust-to-weight ratio of the spacecraft with these engines. Therefore, it makes sense to use electric propulsion engines only in spacecraft after reaching the 1st escape velocity using a chemical or nuclear thruster (in addition, some electric propulsion engines can generally only operate in the vacuum of space).

The idea of using electrical energy to produce jet thrust was discussed by K. E. Tsiolkovsky and other pioneers of astronautics. In 1916-17, R. Goddard confirmed the reality of this idea with experiments. In 1929-33, V. P. Glushko created an experimental electrothermal RD. Then, due to the lack of means of delivering electric propulsion engines into space and the difficulty of creating power supplies with acceptable parameters, the development of electric propulsion engines was stopped. They resumed in the late 50s and early 60s. and were stimulated by the successes of astronautics and high-temperature plasma physics (developed in connection with the problem of controlled thermonuclear fusion). By the beginning of the 80s. In the USSR and the USA, about 50 different designs of electric propulsion systems were tested as part of spacecraft and high-altitude atmospheric probes. In 1964, electromagnetic (USSR) and electrostatic (USA) thrusters were tested for the first time in flight; in 1965, electrothermal thrusters (USA) were tested. Electric propulsion engines were used to control the position and correction of spacecraft orbits, to transfer spacecraft to other orbits (for more details, see the article on various types of electric propulsion engines). Significant progress in the creation of electric propulsion engines has been achieved in Great Britain, Germany, France, Japan, and Italy. Design studies have shown the feasibility of using electric propulsion engines in spacecraft jet control systems designed for long-term operation (several years), as well as as propulsion engines for spacecraft performing complex near-Earth orbital transitions and interplanetary flights. The use of electric propulsion engines instead of chemical thrusters for these purposes will increase the relative mass of the spacecraft payload, and in some cases reduce flight time or save money.

Due to the low acceleration imparted to the spacecraft by electric engines, propulsion systems with electric propulsion propulsion must operate continuously for several months (for example, when a spacecraft is transferring from a low orbit to a geosynchronous one) or several years (during interplanetary flights). In the USA, for example, a propulsion propulsion system with several ion electric propulsion engines with a thrust of 135 mN and a specific impulse of ~ 30 km/s, powered by a solar power plant, was studied. Depending on the number of electric propulsion and the reserve of RT (mercury), the propulsion system could ensure the flight of a spacecraft to comets and asteroids, the launch of a spacecraft into the orbits of Mercury, Venus, Saturn, Jupiter, the sending of a spacecraft capable of delivering Martian soil to Earth, the sending of research probes into the outer atmospheres planets and their satellites, launching spacecraft into circumsolar orbits outside the ecliptic plane, etc. In particular, a propulsion system in the version with 6 electric propulsion engines and a RT reserve of 530 kg could provide a flyby near the comet Encke-Backlund of a payload weighing 410 kg (including 60 kg of scientific equipment).

PSs with electric propulsion engines powered by nuclear power plants are also being studied. The use of these installations, the parameters of which do not depend on external conditions, seems appropriate when the electrical power of the spacecraft is over 100 kW. The indicated propulsion systems can provide maneuvers of transport ships near the Earth, as well as flights between the Earth and the Moon, sending spacecraft for a detailed study of the outer planets, flights of interplanetary manned spacecraft, etc. According to preliminary studies, a spacecraft with an initial mass of 20-30 tons, equipped with a reactor a power plant with a power of several hundred kW and a small number of pulsed electromagnetic electric propulsion engines with a thrust of several tens of N, could study in detail the Jupiter system within 8-9 years, delivering soil samples of its satellites to Earth. Achieving high design characteristics of the propulsion system for such a spacecraft requires, however, solving many problems.

The development of electric propulsion engines contributes to the solution of theoretical issues and the creation of special materials, technology, processes, elements and devices that are of great importance for the development of industrial technological processes, electrical engineering, electronics, laser technology, thermonuclear physics, gas dynamics, as well as space, chemical and medical research.