Electromagnetic rocket engine operating principle. Conversations about rocket engines. Advantages of liquid RDs
In this case, two indicators are identified that reflect the cost of full power when servicing the consumer. These indicators are called active and reactive energy. Total power is the sum of these two indicators. We will try to talk about what active and reactive electricity is and how to check the amount of accrued payments in this article.
Full power
According to established practice, consumers do not pay for the useful power, which is directly used in the household, but for the full power, which is supplied by the supplier. These indicators are distinguished by units of measurement - total power is measured in volt-amperes (VA), and useful power - in kilowatts. Active and reactive electricity is used by all electrical appliances powered from the network.
Active electricity
The active component of total power performs useful work and is converted into those types of energy that the consumer needs. For some household and industrial electrical appliances, the active and apparent power coincide in the calculations. Among such devices are electric stoves, incandescent lamps, electric ovens, heaters, irons, etc.
If the passport indicates an active power of 1 kW, then the total power of such a device will be 1 kVA.
Reactive electricity concept
This is inherent in circuits that contain reactive elements. Reactive electricity is that part of the total incoming power that is not spent on useful work.
In DC circuits there is no concept of reactive power. In circuits, a reactive component occurs only when an inductive or capacitive load is present. In this case, there is a mismatch between the phase of the current and the phase of the voltage. This phase shift between voltage and current is indicated by the symbol “φ”.
With an inductive load, a phase lag is observed in the circuit; with a capacitive load, the phase is advanced. Therefore, only part of the total power reaches the consumer, and the main losses occur due to useless heating of devices and instruments during operation.
Power losses occur due to the presence of inductive coils and capacitors in electrical devices. Because of them, electricity accumulates in the circuit for some time. After this, the stored energy is fed back into the circuit. Devices that contain a reactive component of electricity include portable power tools, electric motors and various household appliances. This value is calculated taking into account a special power factor, which is designated as cos φ.
Reactive power calculation
The power factor ranges from 0.5 to 0.9; The exact value of this parameter can be found in the electrical device data sheet. The apparent power must be determined as the active power divided by the factor.
For example, if the passport of an electric drill indicates a power of 600 W and a value of 0.6, then the total power consumed by the device will be equal to 600/06, that is, 1000 VA. In the absence of passports for calculating the total power of the device, the coefficient can be taken equal to 0.7.
Since one of the main tasks of existing power supply systems is to deliver useful power to the end user, reactive power losses are considered a negative factor, and an increase in this indicator calls into question the efficiency of the electrical circuit as a whole. The balance of active and reactive power in a circuit can be visualized in the form of this funny picture:
The value of the coefficient when taking losses into account
The higher the power factor value, the lower the losses of active electricity will be - which means that the consumed electrical energy will cost the end consumer a little less. In order to increase the value of this coefficient, various techniques are used in electrical engineering to compensate for non-target losses of electricity. Compensating devices are leading current generators that smooth out the phase angle between current and voltage. Capacitor banks are sometimes used for the same purpose. They are connected in parallel to the operating circuit and are used as synchronous compensators.
Calculation of electricity costs for private clients
For individual use, active and reactive electricity are not separated in bills - on the scale of consumption, the share of reactive energy is small. Therefore, private customers with power consumption up to 63 A pay one bill, in which all consumed electricity is considered active. Additional losses in the circuit for reactive electricity are not separately allocated and are not paid for.
Reactive electricity metering for enterprises
Another thing is enterprises and organizations. A huge number of electrical equipment are installed in production facilities and industrial workshops, and the total supplied electricity contains a significant portion of reactive energy, which is necessary for the operation of power supplies and electric motors. Active and reactive electricity supplied to enterprises and organizations requires a clear separation and a different method of payment for it. In this case, the basis for regulating relations between the electricity supply company and end consumers is a standard contract. According to the rules established in this document, organizations that consume electricity above 63 A need a special device that provides reactive energy readings for accounting and payment.
The network company installs a reactive electricity meter and charges according to its readings.
Reactive Energy Factor
As mentioned earlier, active and reactive electricity are highlighted in separate lines in payment invoices. If the ratio of the volumes of reactive and consumed electricity does not exceed the established norm, then no charge for reactive energy is charged. The ratio coefficient can be written in different ways, its average value is 0.15. If this threshold value is exceeded, the consumer enterprise is recommended to install compensating devices.
Reactive energy in apartment buildings
A typical consumer of electricity is an apartment building with a main fuse, consuming electricity in excess of 63 A. If such a building contains exclusively residential premises, there is no charge for reactive electricity. Thus, residents of an apartment building see in the charges payment only for the total electricity supplied to the house by the supplier. The same rule applies to housing cooperatives.
Special cases of reactive power metering
There are cases when a multi-storey building contains both commercial organizations and apartments. The supply of electricity to such houses is regulated by separate Acts. For example, the division can be the size of the usable area. If in an apartment building commercial organizations occupy less than half of the usable space, then reactive energy payments are not charged. If the threshold percentage has been exceeded, then obligations to pay for reactive electricity arise.
In some cases, residential buildings are not exempt from paying for reactive energy. For example, if a building has elevator connection points for apartments, charges for the use of reactive electricity occur separately, only for this equipment. Apartment owners still pay only for active electricity.
Understanding the essence of active and reactive energy makes it possible to correctly calculate the economic effect of installing various compensation devices that reduce losses from reactive loads. According to statistics, such devices allow you to increase the cos φ value from 0.6 to 0.97. Thus, automatic compensating devices help save up to a third of the electricity provided to the consumer. A significant reduction in heat losses increases the service life of devices and mechanisms at production sites and reduces the cost of finished products.
The invention relates to the field of creating electric rocket engines. An electric rocket engine device is proposed, which, just like the known type of engine with a uniform stationary plasma discharge (stationary plasma engines - SPD), contains supersonic nozzles, a magnetohydrodynamic accelerator channel located in a cylindrical cavity between the poles of a coaxial magnetic circuit, a magnetic field excitation coil connected to the EMF source. Unlike the SPD, the proposed engine uses a non-uniform gas-plasma flow of the working fluid. To create plasma inhomogeneities in the form of plasma rings, the engine contains a pulsed high-frequency voltage source connected to an additional coil installed at the input of the accelerator channel. The discharge in the plasma rings, inductively coupled to the magnetic field excitation coil, is maintained by a source of alternating emf connected to the coil. To break the current in the plasma rings at the moment of their exit from the channel of the magnetodynamic accelerator, radial dielectric ribs are installed at the entrance to the engine diffuser. The invention makes it possible to increase the thrust and duration of engine operation. 1 ill.
The invention relates to the field of creating electric rocket engines. There is a known method [I], which increases the thrust of an electric rocket engine, which proposes replacing a stationary homogeneous plasma discharge with a non-uniform gas-plasma flow. Plasma bunches (T-layers) are resistant to the development of overheating instability, which makes it possible to repeatedly increase the density of the working fluid passing through the engine channel, and thus proportionally increase thrust. The device that implements this method consists of a gas-dynamic nozzle, a magnetohydrodynamic accelerator channel of rectangular cross-section with electrode walls, a magnetic system that creates a magnetic field in the accelerator channel transverse to the flow of the working fluid, a pulsed electrode high-current discharge system that forms T-layers in the flow, a source constant EMF connected to the electrodes of the accelerator channel. The device must provide flow acceleration due to the electrodynamic force acting in the volume of T-layers, which in turn act on the gas flow as accelerating plasma pistons. Numerical modeling of the operating mode in the channel of this device has shown that an exhaust velocity of up to 50,000 m/s can be achieved at a thrust level of up to 1000 N. A disadvantage of a device that implements the known method is the use of electrodes both in the source circuit that forms the T-layers, and in source circuit providing acceleration mode in the MHD channel. The mode of current flow in T-layers is arc. The inevitable arc erosion of the electrodes significantly reduces the service life of the engine (from experience with plasma torches, it should be expected that the electrodes will provide no more than 100 hours of continuous operation). For reusable spacecraft, the engine life must be at least a year of continuous operation. An electric rocket engine (stationary plasma engine - SPD) is known, which is used to accelerate the plasma flow due to the electrodynamic effect on the electrically conductive medium. This device consists of supersonic nozzles, a magnetohydrodynamic (MHD) accelerator channel located in a cylindrical cavity between the poles of a coaxial magnetic circuit, a magnetic field excitation coil connected to a constant EMF source, and a power supply system for a stationary plasma discharge. The device operates according to the following scheme. A gaseous working fluid is supplied through a gas-dynamic nozzle, which, upon entering the channel of the MHD accelerator, enters the region of a stationary plasma discharge supported by the power supply system, ionizes and passes into the plasma state. The current in the discharge flows along the channel, while the anode of the power supply system is a gas-dynamic nozzle, and the cathode is located at the outlet of the channel. A stable acceleration mode is realized only at a very low plasma density, at which the Hall parameter can reach values of the order of 100. Under these conditions, a small discharge current along the channel generates a significant azimuthal current, closed on itself. The interaction of the azimuthal current with the radial magnetic field created by the excitation coil between the coaxial poles of the magnetic circuit generates an accelerating electrodynamic force in the plasma volume. The closure of the main current without the use of electrodes for this makes it possible to make the operating life of the engine almost unlimited. The disadvantage of the known device is the low density of the working fluid, which is necessary to ensure stable operation of the engine. Accordingly, the thrust of such an engine does not exceed 0.1 N. The invention is based on the task of creating a high-thrust electric rocket engine with a duration of continuous operation of the order of a year. This task is achieved by the fact that an electric rocket engine containing supersonic nozzles, a magnetohydrodynamic accelerator channel located in a cylindrical cavity between the poles of the coaxial magnetic circuit, the magnetic field excitation coil connected to the EMF source, according to this invention, is equipped with a pulsed high-frequency voltage source connected to an additional coil installed at the input of the accelerator channel, and a diffuser with radial dielectric ribs, while the magnetic field excitation coil is connected to a source of variable EMF. The invention is illustrated by a drawing showing a cross section of the device. An electric rocket engine contains supersonic nozzles 1, channel 2 of a magnetohydrodynamic accelerator located in a cylindrical cavity between the poles of a coaxial magnetic circuit 3, a magnetic field excitation coil 4 connected to a source 5 of a variable EMF, pulsed high-frequency voltage source 6, connected to an additional coil 7 installed at the input to channel 2 of the accelerator. The engine also contains a diffuser 8 with radial dielectric fins 9. An electric rocket engine operates as follows. Heated gas (for example, hydrogen), the temperature of which is determined by the conditions of the on-board heat source, and the pressure is determined by the requirements for engine thrust, which set the flow rate of the working fluid, is accelerated in supersonic nozzle 1. The pulsed high-frequency discharge system 6 is periodically turned on with a given time duty cycle, and each turn on forms a plasma clot in the gas flow at the input of channel 2 of the MHD accelerator. An external source of alternating EMF creates an alternating current in the excitation coil 4, which generates a time-varying radial magnetic field between the poles of the coaxial magnetic circuit 3. This generates an eddy electric field in the azimuthal direction. Under the influence of azimuthal electric and radial magnetic fields, self-sustaining azimuthal plasma current coils (T-layers) are formed from plasma clots, which in turn act on the gas flow as accelerating pistons. After the channel of the MHD accelerator, the accelerated flow enters the expanding channel-diffuser 8, in which radial dielectric fins 9 are installed. The fins are flown around by a gas flow, but the electrical circuits of the T-layers are broken on them, which makes it possible to interrupt the electrodynamic stage of flow acceleration. In diffuser 8, which is a continuation of the channel of the MHD accelerator, further acceleration of the gas flow is carried out due to thermal energy transferred from the T-layers to the flow. Numerical modeling of the process of accelerating the flow of hydrogen containing T-layers was carried out under conditions of a mode that implements the described method . It is shown that the proposed device can be implemented with the following parameters corresponding to the task of creating an efficient electric rocket engine (ERE): - the efficiency of the process of transforming electricity into the kinetic energy of the working fluid is 95%; - the average flow speed at the engine exit is 40 km/s; - the length of the MHD accelerator channel is 0.3 m; - the average diameter of the MHD accelerator channel is 11 cm; - the height of the channel (distance between poles) is 1 cm; - the mass flow rate of the working fluid is 12 g/s; - the temperature of hydrogen at the entrance to the electric propulsion engine is 1000 K; - hydrogen pressure at the entrance to the electric propulsion engine is 10 4 Pa; - the average value of the emf of the electric propulsion power supply is 5 kV; - the average value of the current in the excitation winding is 2 kA; - the consumed electrical power is 10 MW; - the engine thrust is 500 N. The proposed electric rocket engine will find application in the creation space transport system intended for transporting cargo from near-Earth orbits to geostationary, lunar and further to the planets of the solar system. Sources of information1. B.C. Slavin, V.V. Danilov, M.V. Kraev. A method for accelerating the flow of a working fluid in a rocket engine channel, RF patent No. 2162958, F 02 K 11/00, F 03 H 1/00, 2001. 2. S.D. Grishin, L.V. Leskov. Electric rocket engines of spacecraft. - M.: Mechanical Engineering, 1989, p. 163.
Claim
An electric rocket engine containing supersonic nozzles, a magnetohydrodynamic accelerator channel located in a cylindrical cavity between the poles of a coaxial magnetic circuit, a magnetic field excitation coil connected to an EMF source, characterized in that the device is equipped with a pulsed high-frequency voltage source connected to an additional coil installed at the input accelerator channel, and a diffuser with radial dielectric fins, while the magnetic field excitation coil is connected to a source of alternating EMF.
Similar patents:
The invention relates to plasma technology and can be used in electric rocket engines based on a plasma accelerator with closed electron drift, as well as in technological accelerators used in vacuum plasma technology processes
ELECTRIC ROCKET ENGINES(electric propulsion engines, electric propulsion engines) - space. jet engines, in which the directional movement of the jet stream is created due to electricity. energy. An electric propulsion system (EPS) includes the electric propulsion system itself, a system for supplying and storing the working substance, and a system that converts electrical power. parameters of the electric power source to the nominal values for the electric propulsion engine and control the operation of the electric propulsion engine. Electric propulsion engines are low-thrust engines operating for a long time. time (years) on board the spacecraft. aircraft (SC) in conditions of weightlessness or very low gravity. fields. With the help of electric propulsion, the parameters of the spacecraft's flight path and its orientation in space can be maintained with a high degree of accuracy or changed within a given range. With el-magn. or el-static. during acceleration, the exhaust speed of the jet stream in the electric propulsion engine is significantly higher than in liquid or solid-fuel rocket engines; this gives a gain in the payload of the spacecraft. However, electric propulsion engines require a source of electricity, while in conventional rocket engines the energy carrier is fuel components (fuel and oxidizer). The ERD family includes plasma engines
(PD), el-chem. engines (ECM) and ion engines (ID).
Electrochemical motors. In ECD, electricity is used for heating and chemical. decomposition of the working substance. EHD engines are divided into electric heating (END), thermocatalytic (TCD) and hybrid (HD) engines. In the END, the working substance (hydrogen, ammonia) is heated by an electric heater and then flows at supersonic speed through a nozzle (Fig. 1). In the TCD, a catalyst is heated with electricity (to a temperature of ~500 o C), which chemically decomposes the working substance (ammonia, hydrazine); then the decomposition products flow out through the nozzle. In the gas turbine, the working substance is first decomposed, then the decomposition products are heated and flow out. ECD design and used structures. materials are designed to be switched on board the spacecraft for 7-10 years with a number of launches of up to 10 5 , a duration of continuous operation of ~ 10-100 hours and a deviation of thrust characteristics from the nominal value of no more than 5-10%. Electrical power consumption level power - tens of W, thrust range - 0.01 -10 N. ECMs have very low energy for electric propulsion engines. thrust price ~3 kW/N, high jet speed (3 km/s) due to the low molecular weight of the working substance and its decomposition products. A hydrazine gas engine with a thrust of 0.44 H successfully operated on the Intel-sat-5 communications satellite; an ammonia END with a thrust of 0.15 N is part of the standard electric propulsion system of the Meteor series satellites, which corrects the orbit and orientation of the satellite.
Rice. 1. Electric heating motor circuit: 1 - porous electric heater; 2-heat shield; 3 - casing; 4- nozzle.
Ion engines. Will put it in the ID. ions of the working substance are accelerated into electric static. field. ID (Fig. 2) consists of an ion emitter 4, an accelerating electrode 5 with holes (slots) through which accelerated ions pass, and an external electrode. electrode 6 (screen), in the role of which the ID housing is usually used. The accelerating electrode is in negative. potential (~10 3 -10 4 V) relative to the emitter. Electric current and spaces. electric The jet stream must be zero, so the emerging ion beam is neutralized by electrons, which are emitted by neutralizer 7. Ext. the electrode is at a potential negative relative to the emitter and positive relative to the accelerating electrode; positive The potential shift is chosen such that relatively low-energy electrons from the neutralizer are electrically blocked. field and did not fall into the accelerating gap between the emitter and the accelerating electrode. The energy of accelerated ions is determined by the potential difference between the emitter and the external one. electrode. Availability is positive. spaces. charge in the accelerating gap limits the ion current from the emitter. Basic ID parameters: exhaust speed, traction efficiency, energy. thrust price (W/N), energy. ion price (eV/ion) - the amount of energy spent on the formation of an ion. The degree of the working substance in ID should be as high as possible (>0.90.95).
Rice. 2. Diagram of an ion engine with volumetric ionization designs by G. Kaufman: 1 - gas-discharge chamber cathodery; 2- anode; 3 - magnetic coil; 4-emitting electrode; 5 - accelerating electrode; 6 - external electrode; 7 - neutralizer.
Depending on the type of emitter, IDs are divided into surface ionization engines (SSI), colloidal engines (CD) and volume ionization engines (VID). In IDPI, ionization occurs when vapors of the working substance are passed through a porous emitter; the working substance must be less than the work function of the emitter material. Usually a pair of cesium (working substance) - tungsten (emitter) is selected. The emitter is heated to a temperature of 1500 o K to avoid condensation of the working substance. In CD (only laboratory prototypes exist), the working substance (20% solution of potassium iodide in glycerol) is sprayed through capillaries in the form of positively charged microdrops into the accelerating gap; electric the charge of microdroplets arises during the extraction of streams from capillaries in a strong electric current. field and their subsequent disintegration into droplets. The source of ions in the IDP is a gas discharge chamber (GDC), in which atoms of the working substance (metal vapors, inert gases) are ionized by electron impact in a low-pressure gas discharge [discharge between electrodes 1 and 2 (Fig. 2) or electrodeless microwave discharge ]; ions from the GRK are drawn into the accelerating gap through the holes of the emitting electrode-wall of the GRK, which together with the accelerating electrode forms an ion-optical. system (IOS) for accelerating and focusing ions. The walls of the GRK, except for the emitting electrode, are magnetically insulated from the plasma. IDOI - max. developed with engineering and physical From the point of view of IDs, their traction efficiency is ~70%, the operating life confirmed in ground tests is increased to 2 10 4 hours. The operating life of IDs is limited by erosion of the accelerating electrode due to its cathode sputtering by secondary ions resulting from the recharging of fast accelerated ions on slow neutral atoms working substance. Energy the thrust and ion prices in ID (with the exception of CD) are very significant (2·10 4 W/H, 250 eV/ion). For this reason, thrusters are not yet used in space as working electric propulsion engines (ECD, PD), although they have been repeatedly tested on board spacecraft. Naib. significant test under the SERT-2 program (1970, USA); The electric propulsion system included two IDPs designed by G. Kaufman (working fluid - mercury, power consumption 860 W, efficiency 68%, thrust 0.03 H), which worked continuously without failure for 3800 hours and 2011 hours, respectively, and resumed operation after a long period. break.
PD according to the scheme of plasma accelerators with closed electron drift and an extended acceleration zone is systematically used on spacecraft, especially on geostationary communications satellites.
Lit.: Gilzin K. A., Electric interplanetary ships, 2nd ed., M., 1970; Morozov A.I., Shubin A.P., Space electric propulsion engines, M., 1975; Grishin S. D., Leskov L. V., Kozlov N. P., Electric rocket engines, M., 1975.
Electric rocket motor (ERD)
The limited use of electric propulsion engines is associated with the need for high power consumption (10-100 kW by 1 n traction). Due to the presence of an on-board power plant (and other auxiliary systems), as well as due to the low thrust density, a device with an electric propulsion engine has low acceleration. Therefore, electric propulsion engines can only be used in spacecraft flying either in conditions of weak gravitational fields or in near-planetary orbits. They are used for orientation, correction of spacecraft orbits and other operations that do not require large amounts of energy. Electrostatic, plasma Hall and other electric propulsion systems are considered promising as the main engines of spacecraft. Due to the small ejected mass of the RT, the continuous operation time of such electric propulsion engines will be measured in months and years; their use instead of existing chemical taxiways will increase the payload mass of the spacecraft. The idea of using electrical energy to generate thrust was put forward by K. E. Tsiolkovsky and other pioneers of astronautics. In 1916-17, R. Goddard (USA) confirmed the reality of this idea with experiments. In 1929-33, V. P. Glushko (USSR) created an experimental electric propulsion engine. In 1964 in the USSR, plasma pulsed thrusters were tested on the Zond-type spacecraft, in 1966-71 on the Yantar spacecraft - ion thrusters, in 1972 on the Meteor spacecraft - plasma quasi-stationary thrusters. Various types of electric propulsion systems have been tested since 1964 in the USA: in ballistic flight, and then in space flight (on ATS, CERT-2, etc.). Work in this area is also being carried out in Great Britain, France, Germany, and Japan. Lit.: Corliss W.R., Rocket engines for space flights, trans. from English, M., 1962; Stuhlinger E., Ion engines for space flights, trans. from English. M., 1966; Gilzin K. A., Electric interplanetary ships, 2nd ed., M., 1970; Gurov A.F., Sevruk D.D., Surnov D.N., Design and strength calculation of space electric rocket engines, M., 1970; Favorsky O. N., Fishgoit V, V., Yantovsky E. I., Fundamentals of the theory of space electric propulsion systems, M., 1970; Grishin S. D., Leskov L. V., Kozlov N. P., Electric rocket engines, M., 1975. Yu. M. Trushin.
Great Soviet Encyclopedia. - M.: Soviet Encyclopedia. 1969-1978 .
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