CN104696180A - Magnetic field regulation type large-area microcavity discharge plasma micro-thruster for liquid working substances - Google Patents

Abstract

The magnetic field control type liquid-phase working medium large-area microcavity discharge plasma micro-thruster, the propellant working medium feeding mechanism diffuses the liquid propellant in the liquid propellant storage tank into the discharge microcavity, and the discharge microcavity is composed of A hole with different diameters runs through the metal sheet and the insulating dielectric, and the insulating dielectric is sandwiched between two adjacent metal sheets, and the adjacent two metal sheets form an electrode pair; the metal sheet at the end of the discharge microcavity and the insulating dielectric form a pull Farr nozzle, where the insulating dielectric forms the constriction and divergence of the Rafal nozzle, and the anode metal sheet forms the throat of the Rafal nozzle, which forms an electrode with the cathode metal sheet located at the exit of the divergent section of the Rafal nozzle Yes; at least one electrode pair is set on the horizontal section of the Rafal nozzle, the external power supply module provides periodic AC or DC high-voltage pulse voltage, and is connected to each electrode pair with a wire, and the ring-shaped permanent magnet is set in the above discharge microcavity outside of the pipe wall.

Description

Large-area microcavity discharge plasma microthruster with magnetic field control type liquid working medium

technical field

The invention relates to a large-area micro-cavity discharge plasma space micro-propulsion based on magnetic field modulation, which is especially suitable for micro-propulsion systems with high integration, low power consumption, precise impulse, and high specific impulse for space micro-spacecraft work needs.

Background technique

The urgent need for micro-propulsion technology for space miniaturized platforms stems from the vigorous development of micro-nano satellites in this century. After entering the 21st century, micro-nano satellite networking technology has developed rapidly and become increasingly mature, which enables micro-nano satellites to perform the functions of traditional large satellites while having the advantages of short cycle, low risk, flexible launch methods, and strong scalability . Due to the small size, light weight and small moment of inertia of the micro-nano satellite, the thrust required for satellite orbit and attitude control is small, the repeatability is good, and the precision is required to be high, generally in the order of millinewtons, or even in the order of micronewtons; excellent Micro propulsion should also have a wide range of thrust to meet the needs of different mission applications. At the same time, the goal of micro-nano-satellite formation flight is to maintain the relative position between satellites, not to maintain the absolute position of each satellite, so the minimum impulse pulse required is very small to meet the high-precision requirements of the control system. Therefore, research on micro-propulsion systems with high integration, low power consumption, precise impulse, and high specific impulse suitable for micro-nano-satellite orbit keeping, orbital maneuvering, and attitude control is an inevitable requirement for the development of micro-nano-satellites and a key link to success or failure.

Compared with chemical micro-propulsion, electric propulsion has the advantages of low cost, high specific impulse, small thrust, and high control precision. It can greatly reduce the amount of propellant carried and increase the payload of the satellite, or greatly extend the working life of the satellite without reducing the propellant. , thus becoming one of the most potential micro-propulsion methods. Micro-electric propulsion is divided into electrothermal, electrostatic and electromagnetic. Electrothermal micro-propulsion is non-polluting, in principle, any propellant can be used, and the cost is low, but its specific impulse is very low, and there is a leakage problem. Electrostatic micro-thrusters include field-effect electrostatic thrusters, colloidal micro-thrusters, ion micro-thrusters, and Hall micro-thrusters. However, electrostatic micro-propellers generally require high operating voltage (kV order) or relatively large power supply (hundred-watt order), which exceeds the supply capacity of micro-nano satellites, which limits its application in ultra-small micro-nano satellites. Applications. Electromagnetic micro-thrusters such as pulsed plasma thrusters have a large inherent mass, also have the problem of plume pollution, and the thrust power ratio is not high.

So far, the constraints of mass and power on the propulsion system have made the electric propulsion method of micro-nano satellites mainly electrothermal. The traditional resistance heating micro-propulsion system faces outstanding problems such as small specific impulse and low heating efficiency. The radio-frequency discharge electrothermal micro-propulsion has the disadvantages of low energy utilization efficiency and large ablation of the wall surface, and the size of the micro-arc engine cannot meet the requirements of centimeters processing needs. These unfavorable factors limit the application of the above-mentioned propulsion system on micro-nano satellites. In contrast, microcavity discharge plasma propulsion is a new electrothermal micro-propulsion technology developed from micro-discharge technology. Microcavity discharge is a novel non-equilibrium high-pressure glow discharge, which only requires very low voltage (hundreds of volts) or input power (on the order of hundreds of milliwatts), through the collision of plasma and neutral gas and the electric field It is used for the acceleration of ions, heats the working fluid gas, and then sprays out through a certain nozzle structure to generate thrust, thus opening up a new way to improve the performance of the propulsion system. The Electric Propulsion Laboratory of Illinois State University and the Computational Plasma Laboratory of Texas State University first proposed the design of the micro-cavity discharge micro-nano satellite thruster (as shown in Figure 1) and received the US Air Force Center Scientific Research Fund The corresponding research work was carried out under the support of <R.L.Burton et al. Development of the MCD thruster for nanosat propulsion, Proceedings of the 57th Joint Army Navy NASA Air Force Propulsion Meeting, Colorado Springs, Colorado, 2010; M.de Chadenedes et al. thruster technology, Proceedings of the 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Nashville, TN, 2010>. At present, there are many outstanding problems in the micro-cavity discharge plasma thruster, which restricts the further improvement of thrust, specific impulse and energy efficiency: (1) Due to the large surface area to volume ratio of the micro-discharge cavity, a large amount of power loss occurs In the sheath region on the interface between the solid wall and the plasma, the energy utilization efficiency of the heating and propulsion of the propellant working fluid is seriously reduced. (2) Due to the strong coupling between the plasma and the cold propellant gas flow, the charged ions are mainly distributed inside the discharge cavity or nozzle, and the real thrust contributed by the charged particles is very limited. (3) The discharge area of the single-pair electrode structure is small, and the discharge stability is not high. When the propellant flow rate is large, the ionization efficiency is limited, and a large amount of gas-phase working fluid cannot be effectively discharged, ionized and heated, resulting in low working medium utilization efficiency of the propeller. (4) Gas propellants have high requirements on mass flow control accuracy and require high-pressure storage, which will occupy the volume and mass of tiny spacecraft. Based on the above problems, the specific impulse of the current microcavity discharge plasma micro-thruster is only 100-200s, the energy efficiency is less than 60%, and the propulsion control accuracy is not high.

Contents of the invention

The technical problem of the present invention is: to overcome the bottlenecks of traditional micro-cavity discharge plasma micro-thrusters such as low specific impulse, low energy efficiency, and low propulsion control precision, and propose a large-area micro-cavity discharge plasma based on magnetic field modulation Body micro propulsion, through the comprehensive regulation of various external control conditions and the optimal configuration of various new components to achieve high specific impulse, long life and high control precision micro-nano satellite propulsion.

The technical solution of the present invention is: a magnetic field control type liquid-phase working medium large-area microcavity discharge plasma micro-thruster, including a thruster body, a permanent magnet and an external power supply module; the thruster body structure includes a liquid propellant storage tank, propelling Agent feeding mechanism, sheet metal, dielectric and Rafal nozzle;

The propellant working medium feed mechanism controls the flow of the working medium and forms a spray of liquid propellant in the liquid propellant storage tank to spread into the discharge microcavity. The discharge microcavity is formed by a hole with different diameters that penetrates the metal sheet and the insulation. The insulating dielectric is sandwiched between two adjacent metal sheets, and the adjacent two metal sheets form an electrode pair; the metal sheet and the insulating dielectric at the end of the discharge microcavity form a Rafah spout, and the insulating dielectric constitutes a Rafah The shrinkage section and the expansion section of the Rafal nozzle, the throat of the anode metal sheet constitutes the Rafal nozzle, and it forms an electrode pair with the cathode metal sheet located at the outlet of the Rafal nozzle expansion section; on the horizontal section of the Rafal nozzle, at least An electrode pair is set, the external power supply module provides periodic AC or DC high-voltage pulse voltage, and is connected to each electrode pair by a wire, and the ring-shaped permanent magnet is set on the outside of the tube wall of the above-mentioned discharge microcavity.

The above-mentioned external power supply module and propeller body structure are used as a propulsion unit, and multiple identical propulsion units are arranged in a lattice, and the external ring-shaped permanent magnets form a multi-cavity propeller, and insulating materials are filled between the permanent magnets and each propulsion unit .

Compared with the prior art, the present invention has beneficial effects as follows:

(1) The surface discharge of the compound multi-electrode structure is used to obtain a large-area discharge plasma at a low power input power surface density, and can reduce the high-frequency oscillation of the discharge, improve the stability of the discharge, and overcome the large flow rate of the propellant A large amount of gas-phase working fluid cannot be effectively discharged, ionized and heated, and the bottleneck of low utilization efficiency of the thruster working medium;

(1) The external axial magnetic field modulation optimization generated by the circular permanent magnet can significantly reduce the discharge voltage and strengthen the ionization process, regulate the transport characteristics of the microcavity discharge plasma discharge, and improve the contribution of charged particles to the improvement of propulsion performance;

(2) The diffusion speed of liquid propellant into the discharge microcavity through the working fluid feeding mechanism is much slower than that of gas propellant, and the flow control of propellant is easier, which can realize the specific impulse, thrust and micro impulse of micro-nano satellite thrusters Higher-precision operation control; secondly, the liquid propellant has a high mass density and does not require a high-pressure storage tank, which can save the volume and weight of the propeller and the micro-nano satellite;

(3) The micro-cavity discharge plasma micro-thruster is manufactured with MEMS technology to meet the stringent requirements of micro-nano satellites for propulsion system weight, volume, and power;

(4) The working medium feeding mechanism makes the liquid propellant enter the discharge microcavity in the form of mist, thereby increasing the contact area between the liquid propellant, the discharge electrode and the plasma, which is more conducive to the heating of the propellant and the improvement of the discharge efficiency .

(5) Multiple thruster modules of the same type are integrated and combined to form a multi-cavity thruster, which can realize thrust adjustment in a large range, and is applied to propulsion tasks with different thrust requirements, and has the capability of continuously adjustable thrust.

Description of drawings

Fig.1 Schematic diagram of traditional microcavity discharge thruster

Fig. 2 is a schematic diagram of a single propulsion unit of the present invention;

Figures 3a and 3b are two schematic diagrams of the multi-cavity propeller arrangement of the present invention.

Detailed ways

As shown in Figure 2, a magnetic field control type liquid-phase working medium large-area microcavity discharge plasma micro-thruster of the present invention includes a thruster body structure, a permanent magnet 9 and an external power module 6; the thruster body structure includes a liquid propulsion Agent storage tank 8, propellant working medium feeding mechanism 7, metal sheet (anode 1, cathode 2), dielectric 3 and Rafal nozzle 4;

Wherein, the liquid propellant storage tank 8 stores volatile liquid propellants (ethanol, water, etc.), and the working medium feed mechanism 7 includes a micro valve for controlling the flow of the working medium and a capillary injection for atomizing the liquid working medium. Atomizer, which diffuses the liquid propellant into the discharge microcavity and realizes spray treatment. In the space environment, the capillary injection atomizer presses the liquid working medium into the capillary to form a high-speed jet, and ejects the liquid droplets as extremely fine particles to form a liquid mist, which can automatically diffuse into the electrode space where the outlet end is a vacuum. The discharge microcavity is composed of a hole with a diameter (from tens of microns to hundreds of microns) that runs through a metal sheet (copper, nickel, platinum, tungsten, etc.) and an insulating dielectric 3 (mica, ceramics, organic glass, etc.) , the insulating dielectric 3 is sandwiched between two adjacent metal sheets, and the adjacent two metal sheets form an electrode pair; the dielectric 3 is used to prevent the discharge from occurring in other parts between the electrodes in the electrode chamber.

The above two metal sheets and the insulating dielectric 3 sandwiched between the two metal sheets are called a discharge device. Multiple discharge devices inside the thruster body are placed in parallel, and the metal sheet and the insulating dielectric at the end of the discharge microcavity form a Rafael. Nozzle 4, wherein the insulating dielectric constitutes the constriction and expansion of the Rafal nozzle 4, and the anode metal sheet forms the throat of the Rafal nozzle 4, which is formed with the cathode metal sheet at the exit of the expansion section of the Rafal nozzle 4 One electrode pair; at least one electrode pair is arranged on the horizontal section of the Rafal nozzle 4 .

The external power supply module 6 converts the low-voltage stable DC output obtained by the conversion of the spacecraft solar cell into a periodic AC or DC high-voltage pulse voltage, and connects it to the terminals of the anode 1 and the cathode 2 by wires, and connects the two metal sheet electrodes The plasma 5 is generated by the discharge between them, and the discharge condition is related to the size of the aperture, the distance between the electrodes, the air pressure and the magnitude of the current. The external power sources connected to each electrode pair are connected in parallel, and the discharge is controlled by an external trigger circuit with the same amplitude and frequency to generate plasma. Due to the small aperture of the discharge microcavity, only a low voltage is needed to maintain a stable discharge between the two electrodes, and a plasma with a high gas temperature and electron density is obtained, and the unit is heated by the discharge of multiple electrodes. , passing through the end of the Lafal nozzle (with a typical constriction section, confinement section and expansion section), it presents a high-speed flow after expansion in the downstream expansion section, and the propulsion of the spacecraft is achieved with the help of effective reaction force. A plurality of discharge units are arranged in parallel to form a propulsion array to realize automatic adjustment and control of the thrust. The ring-shaped permanent magnet 9 (the magnet is a neodymium-iron-boron permanent magnet, can also be a samarium-cobalt magnet, an alnico magnet, a ferrite magnet, etc.) ) installed on the outside of the tube wall of the above-mentioned discharge microcavity to generate an axial magnetic field consistent with the flow direction of the gas flow, significantly reduce the discharge voltage and strengthen the ionization process, and regulate the movement of charged particles along the parallel axis, so as to modulate the discharge process and plasma output purpose of the shipping process.

As shown in Figure 3, the above-mentioned external power supply module 6 and the thruster body structure are used as a propulsion unit 10, and a plurality of identical propulsion units are arranged in a lattice. In order to facilitate the control of the work of the propulsion unit, the possible arrangement of the propulsion units includes Torus and two-dimensional matrix. The outer ring-shaped permanent magnet 9 forms a multi-cavity propeller, and insulating material is filled between the permanent magnet 9 and each propulsion unit.

Parts not described in detail in the present invention belong to the common knowledge of those skilled in the art.

Claims (2)

1. The magnetic field control type liquid-phase working medium large-area microcavity discharge plasma micro-thruster is characterized in that: it includes a thruster body structure, a permanent magnet (9) and an external power supply module (6); the thruster body structure includes a liquid propulsion agent storage tank (8), propellant working fluid feeding mechanism (7), metal sheet, dielectric (3) and Rafal nozzle (4); wherein, under DC discharge conditions, the metal sheets respectively form a pair and exist anode (1), cathode (2); The propellant working fluid feed mechanism (7) controls the flow of the working fluid and diffuses the liquid propellant in the liquid propellant storage tank (8) into the discharge microcavity, and the discharge microcavity is penetrated by a hole with different diameters The metal sheet and the insulating dielectric (3), the insulating dielectric is sandwiched between two adjacent metal sheets, and the adjacent two metal sheets form an electrode pair; the metal sheet at the end of the discharge microcavity and the insulating dielectric form a Rafal spout ( 4), wherein the insulating dielectric constitutes the contraction section and the expansion section of the Rafal spout (4), and the anode metal sheet forms the throat of the Rafal spout (4), which is located at the outlet of the expansion section of the Rafal spout (4). The cathode metal sheet at the place forms an electrode pair; at least one electrode pair is set on the horizontal section of the Rafael spout (4), and the external power supply module (6) provides periodic AC or DC high-voltage pulse voltage, and is connected to each On the electrode pair, the ring-shaped permanent magnet (9) is sleeved outside the tube wall of the above-mentioned discharge microcavity. 2. the magnetic field control type liquid-phase working medium large-area microcavity discharge plasma micro-thruster according to claim 1 is characterized in that: the above-mentioned external power supply module (6) and the thruster body structure are used as a propulsion unit (10 ), a plurality of identical propulsion units are arranged in a dot matrix, and an outer ring-shaped permanent magnet (9) forms a multi-cavity propeller, and an insulating material is filled between the permanent magnet (9) and each propulsion unit.