RU2818410C1 - Electrojet propulsion system - Google Patents

Abstract

FIELD: space engineering.

SUBSTANCE: invention relates to electric rocket engines with inductive plasma source using outboard atmosphere as working medium. Electrojet propulsion system comprises a gas-collecting device, a heat-stabilizing chamber, hermetically connected to the gas collecting device and a storage tank, gas-discharge chamber installed coaxially to gas collecting device. Gas-discharge chamber is a hollow cylinder, the walls of which are made of radiotransparent dielectric material. One end of the gas-discharge chamber is connected through gas lines, including shutoff and control valves, with a storage tank, and the other is openly communicated with the environment. Working medium ionization and acceleration system includes an antenna located coaxially to the gas-discharge chamber, a high-frequency generator, matching unit of high-frequency generator and antenna, constants or electromagnets and power supply. Outboard gas collection device is adjustable and is made in the form of a structure of variable cross section and/or variable opening angle. Thermostabilizing chamber is tightly connected to the storage tank by gas mains through a blower and a device which prevents gas backflow.

EFFECT: higher spacecraft maneuverability, possibility of its operation in low orbit of atmosphere celestial bodies, higher efficiency.

2 cl, 1 dwg

Description

The invention relates to space technology, in particular to electric rocket propulsion systems with an inductive plasma source, using an outboard atmosphere as a working fluid and intended to increase the service life and implement controlled movement of flying spacecraft in space in low orbits of celestial bodies with an atmosphere.

An air-jet electrostatic ion engine is known (see US patent No. 758I380B2 dated September 1, 2009, IPC class F03HI/0043). An electrostatic ion engine, which uses atmospheric gas as a working fluid, contains a housing, a gas-discharge chamber formed by a non-conducting outer surface and a conducting inner surface, into which atmospheric gas enters and is ionized by electrons emitted by an internal electrode (electron gun). The engine contains electrically charged electrodes at the ends of the chamber, while the emission one passes atmospheric gas into the chamber, ionizing it, and the ion-optical system, consisting of a pair of electrodes, attracts charged ions and accelerates them. One or more magnets form an external magnetic field. The neutralization unit (cathode-neutralizer) at the rear of the housing equalizes the ion-electronic equilibrium and maintains the ion engine at a neutral electrical potential through the cathode-neutralizer.

The main disadvantage of the above design is ion erosion of the system electrodes, which limits the service life, changing the speed and direction of particle movement during the destruction of the electrodes, and, as a consequence, the parameters of the generated thrust. The electrodes also impose restrictions on the range of wave power transmitted to the plasma due to the limitation of the maximum permissible concentration of charged particles, which leads to a dependence of the engine performance on the structurally achievable potential difference across the electrodes. Another disadvantage is the principle of acceleration of the working fluid, which leads to the spatial separation of charges into ions and electrons, the need to use a cathode-neutralizer and a cathode in the engine chamber, which removes excess electrons from the volume of the gas-discharge chamber.

Also, a significant drawback is the lack of possibility of accumulation of atmospheric gas, which makes the operation of the propulsion system directly dependent on local changes in the density of the outboard gas, which causes instability of the engine operating parameters, such as power consumption and thrust produced, which, in turn, provokes a load on on-board power supply and control systems.

Also, a significant drawback is the lack of thermal stabilization of incoming gas molecules, since atmospheric particles have high speeds (relative to the “spacecraft-particle” system under consideration), and, entering the internal volume of the gas-discharge chamber through the first electrode, will subject it to severe erosion, while partially without undergoing the ionization process.

A direct-flow electric jet engine is known (see patent RU No. 2614906 dated 04/05/2016, class IPC F03H 1/00, IPC H05H 1/54), containing a space charge neutralizer of the ion current, a dielectric transport channel, a cylindrical direct-flow channel at the inlet in which a gas collection device with axial channels is installed, the entrance of which is connected to the inlet chamber. The ion-optical system is located in the outlet of the direct-flow channel and includes an emission electrode, an accelerating electrode and a decelerating electrode connected to a power source. The chamber contains a coaxial spiral inductor with a variable cross-sectional area of the surface of rotation, increasing in the direction from the gas collection device to the electrodes of the ion-optical system.

The disadvantage of this invention is the inability to store and controllably consume atmospheric gas, which makes the operation of an electric jet engine dependent on local changes in the density of atmospheric gas, the characteristics of the engine are variable, and the execution of orbital maneuvers is more complex in terms of control and execution. Also disadvantageous is the spatial separation of charges, which requires compensation through the use of a neutralizer, and the use of eroded electrodes to accelerate ionized particles.

The closest known technical solution is an outboard air engine with a helicon plasma source to support small spacecraft in low Earth orbit (see patent No. RU2703854 CI, dated October 22, 2019, class MPKF03H1/00).

The invention relates to electric rocket propulsion systems, in particular to electric rocket engines with a helicon plasma source using the outboard atmosphere as a working fluid, intended mainly for installation on small aircraft. The engine contains a thermal stabilizing channel, a receiver and a gas-discharge chamber in a single motor path, an RF antenna, an RF generator, a device for matching the load and the RF generator, permanent or electromagnets and at least one power supply.

The disadvantage of this prototype is the inability to regulate the amount of atmospheric gas taken in, which makes control of the propulsion system less flexible and the program for performing the orbital maneuver more complex. The unregulated useful area of the air intake causes an increase in the aerodynamic resistance of the spacecraft as the pressure inside the thermal stabilization channel and receiver increases, associated with a decrease in the mass flow of gas in the engine or, if it is constant, with a decrease in the ambient pressure, which leads to the need to increase the duration of operation of the engine or its power to compensate for the loss of the orbital speed of the device.

Also, a disadvantage of this prototype is the inability to store atmospheric gas, since the oncoming flow of particles in various operating conditions of the spacecraft may not be enough to generate thrust that compensates for the aerodynamic braking of the spacecraft, but the prototype design does not provide for any reserves of the working fluid for orbital maneuvers of the spacecraft. apparatus.

Also, a disadvantage of this prototype is the presence of a single propulsion tract, which imposes significant restrictions on the design of the aircraft, since the use of a propulsion system of such a configuration involves either an increase in drag due to the placement of the propulsion system in a separate gondola, or the use of a coaxial arrangement of the remaining elements of the spacecraft relative to the propulsion system .

The technical effect achieved by the proposed electric propulsion system is to increase the maneuverability of the spacecraft, the possibility of its operation in low orbits of celestial bodies with an atmosphere, and the increase in efficiency caused by the use in the design of the propulsion system of an adjustable gas collection device that effectively collects the outboard atmosphere, a supercharger and a gas storage containers that store and consume the working fluid in a controlled manner, preventing its reverse motion while simultaneously stabilizing and slowing down, leading to an increase in the volume concentration of ionized particles in the gas-discharge chamber, as well as gas lines that ensure industrial feasibility through the use of a free layout of the propulsion system as part of the aircraft by separating components and assemblies of the propulsion system. The need to increase the dimensions of the aircraft by moving the propulsion system or design restrictions on its layout is eliminated. The thrust density and the specific thrust impulse are increased due to the use of an inductive plasma source, which increases the degree of ionization of the working fluid.

The specified technical effect is achieved in an electric propulsion system containing a gas collection device, a heat-stabilizing chamber, hermetically connected to the gas collection device and a storage tank, a gas-discharge chamber installed coaxially to the gas collection device and representing a hollow cylinder, the walls of which are made of radio-transparent dielectric material, one end of which is connected through gas lines, including shut-off and control valves, with a storage tank, and the other openly communicated with the external environment, a system for ionization and acceleration of the working fluid, including an antenna located coaxially with the gas-discharge chamber, a high-frequency generator, a matching unit for the high-frequency generator and antennas, permanent or electromagnets and a power source, characterized in that the outboard gas collection device is adjustable and made in the form of a structure of variable cross-section and/or variable opening angle, the thermostabilizing chamber is hermetically connected to the storage tank through gas lines through a supercharger and a device that prevents gas backflow.

The propulsion tract is made in the form of gas lines, and the components and assemblies of the propulsion system are spaced within the aircraft.

The essence of the invention is illustrated in the drawing, which shows that the low-orbit electric propulsion engine includes:

1 - gas collection device;

2 - thermal stabilizing chamber;

3 - gas line;

4 - supercharger;

5 - storage capacity;

6 - gas discharge chamber;

7 - cooling system;

8 - antenna;

9 - RF generator;

10 - matching device;

11 - permanent or electromagnets;

12 - power supply;

13 - shut-off and control valves.

Structurally, the low-orbit electric jet engine contains an adjustable outboard gas collection device (1) and a thermal stabilizing chamber (2) hermetically connected to it at one end, the other end of which is hermetically connected to the supercharger (4) through gas lines (3). The storage tank (5) is hermetically connected by gas lines (3) in series between the supercharger (4) and the gas-discharge chamber (6), on the outer shell of which a cooling system (7) and a coaxial plasma-forming antenna (8) connected to the RF generator ( 9) through a matching device (10); the magnetic system consists of permanent or electromagnets (11), power is provided by at least one power source (12). Gas lines (3) are equipped with elements of shut-off and control valves (13).

The operation of a low-orbit electric propulsion engine is carried out as follows. The gas collection device (1) collects particles of the rarefied outboard atmosphere and directs them into a heat-stabilizing chamber (2) sealed with the gas collection device (1), in which the particles of atmospheric gas, due to repeated collisions with the inner walls of the heat-stabilizing chamber (2) and with each other, align their energy, slowing down the speed, which ultimately leads to an increase in the volume concentration of ionized particles in the gas-discharge chamber (6). The use of a supercharger (4) in the design of the propulsion system, which extracts rarefied gas from the thermostabilizing chamber (2) through gas lines (3) into a storage tank (5), allows, in addition to the effective collection of gas particles, also to increase the amount of gas collection device (1) passing through a unit area. atmospheric gas by maintaining the required ratio of volumetric particle concentrations between the thermostabilizing chamber (2) and the space in front of the gas collection device (1). To ensure the best energy efficiency, the supercharger (4) can be equipped with a device that turns it on only when a certain concentration of sea gas particles is reached in the thermostabilizing chamber (2). The storage tank (5) is hermetically connected by gas lines (3) in series with the supercharger (4) and the gas-discharge chamber (6), and is equipped with elements of shut-off and control valves (13) that prevent the reverse flow of the working fluid and regulate its mass flow in order to maintain constant values of engine operating parameters, such as thrust and specific impulse. In the gas-discharge chamber (6), made of a radio-transparent dielectric material in the form of a hollow cylinder with an internal diameter that satisfies the Debye criterion for the required operating conditions, particles of outboard gas collected by the outboard air collection device (1) are slowed down and equalized in speed in the heat-stabilizing chamber (2), diverted through gas lines by a supercharger (4) into a storage tank (5), stored and uniformly consumed from the storage tank (5), are ionized under the action of Trivelpiece-Gould waves or oblique Langmuir waves, i.e. under the influence of a spiral electromagnetic field, which is a superposition of fields: a constant external magnetic field created by a magnetic system consisting of permanent or electromagnets (11) and at least one power source (12), and an electromagnetic field created by a plasma-forming antenna (8), when applying to which the RF power through the matching device (10) from the HF generator (9) via the electrical communication line plasma antenna (8) - matching device (10). The matching device (10) of the load (plasma-forming antenna (8) and plasma in the gas-discharge chamber (6)) with the RF generator (9), located on the electrical communication line of the plasma-forming antenna (8) with the RF generator (9), adjusts the load impedance (plasma-forming antenna (8) and plasma in the gas-discharge chamber (6)) to values equal to the impedance of the RF generator (9). Thus, on the electrical connection line of the plasma-forming antenna (8) with the RF generator (9), RF power losses are reduced.

The advantage of the above electric propulsion system is an increase in the maneuverability of the aircraft, the possibility of its operation in low orbits of celestial bodies with an atmosphere, and an increase in efficiency, industrial feasibility and the absence of design restrictions on the layout while maintaining miniaturization of the weight and size parameters of the aircraft, high energy efficiency achieved by increasing the density and values of thrust and specific impulse of the propulsion system.

Thus, the implementation of the invention will lead to the possibility of industrial applicability and operation of a new segment of low-orbit propulsion systems and scientific research operation of undeveloped orbits of celestial bodies with an atmosphere.

Claims (2)

1. An electric propulsion system containing a gas collection device, a thermal stabilizing chamber hermetically connected to the gas collection device and a storage tank, a gas discharge chamber installed coaxially to the gas collection device and representing a hollow cylinder, the walls of which are made of radiotransparent dielectric material, one end of which is connected through gas lines , including shut-off and control valves, with a storage tank, and the other is openly communicated with the external environment, a system for ionization and acceleration of the working fluid, including an antenna located coaxially with the gas-discharge chamber, a high-frequency generator, a matching unit for the high-frequency generator and antenna, permanent or electromagnets and a power source , characterized in that the outboard gas collection device is adjustable and made in the form of a structure of variable cross-section and/or variable opening angle, the thermostabilizing chamber is hermetically connected to the storage tank through gas lines through a supercharger and a device that prevents gas backflow. 2. The propulsion system according to claim 1, characterized in that the propulsion tract is made in the form of gas lines, and the components and assemblies of the propulsion system are spaced apart within the aircraft.