Working fluidless micro thermionic emission device

Abstract

The application provides a working medium-free micro thermionic emission device, which comprises an extraction electrode, an electron emitter, a heating assembly, a shielding cover, an insulating connecting piece and a supporting cylinder; the extraction electrode is arranged outside the supporting cylinder, and the electron emitter and the heating assembly are arranged inside the supporting cylinder; the heating assembly is used for heating the electron emitter so that the electron emitter emits a thermionic emission current; the central axis of the electron emitter corresponds to an extraction hole on the extraction electrode; a gap exists between the emission end face of the electron emitter and the extraction electrode; the two ends of the insulating connecting piece are connected with the outer side wall of the supporting cylinder and the inner side wall of the extraction electrode respectively; the shielding cover is arranged on the circumferential side of the supporting cylinder, and a gap exists between the shielding cover and the supporting cylinder. The electron emission device of the application can be used as an electron source for a tether propulsion system, and can greatly improve the performance of the propulsion system, simplify the system structure, reduce the weight and complete space tasks such as orbit transfer of a spacecraft or de-orbiting at the end of the life of the spacecraft.

Description

Technical Field

[0001] This invention relates to the field of aerospace propulsion technology, specifically to a propellantless micro thermionic emission device, and more particularly to a propellantless micro thermionic emission device for a tethered propulsion system. Background Technology

[0002] In commonly used space propulsion technologies, whether chemical or electric, the reaction force of propellant ejection propulsion is relied upon. With the increasing geometrical amount of space debris and junk, spacecraft must possess end-of-life deorbiting capabilities. These space propulsion mission requirements make propellant load a decisive factor limiting a spacecraft's orbital maneuverability and on-orbit lifespan. Therefore, propellant-free, electro-tethered propulsion technology has become a hot topic of research and development among various countries.

[0003] Space tethered propulsion is a novel propulsion technology that generates thrust by cutting through the space magnetic field using a conductive, flexible tether. This technology uses conductive, flexible tethers to connect two or more spacecraft together. A current of a certain magnitude and direction is passed through the conductive tether, forming a closed loop with the plasma layer in the space environment. The conductive tether then generates a Lorentz force in the Earth's magnetic field, such as... Figure 2 As shown, the Lorentz force can provide thrust to the spacecraft. When the Lorentz force is in the same direction as the system's orbital velocity, it achieves orbital lifting. When the current in the tether is reversed, the resulting Lorentz force is in the opposite direction to the spacecraft's orbital velocity, enabling the spacecraft to descend to a lower orbit and complete its end-of-life deorbiting function.

[0004] During its operation in orbit, the space tethered propulsion system emits low-energy electrons into the space environment through an electron emission device at one end of the tether, which couples with the surrounding plasma. At the same time, an electron collection device at the other end collects electrons from space, thereby forming an electric current in the tether and generating a Lorentz force.

[0005] The magnitude of the Lorentz force conforms to formula (1):

[0006]

[0007] In the formula, F is the Lorentz force, B is the geomagnetic field strength, I is the electron current flowing through the conductive tether, l is the length of the conductive tether, and dl is the length of the infinitesimal element of the conductive tether.

[0008] The strength of the Earth's magnetic field in space is 5 × 10⁻⁶. -5 ~5×10 -7Tesla, with a tether length typically ranging from 1 to 10 km depending on the system design, for a specific spacecraft deorbiting mission, the magnitude of the Lorentz force is directly proportional to the magnitude of the electron current, while the deorbiting time is inversely proportional to the electron current. For example, a satellite with a mass of 500 kg and an orbital altitude of 350 km, if deorbited at the end of its life using a space tether propulsion system with a length of 2 km and an electron current of 20 mA, can generate a thrust of 1 to 2 mN, and it will take approximately one year for the orbit to decrease to an orbital altitude of 200 km to complete the deorbiting mission.

[0009] Currently, most research on space tethered propulsion uses hollow cathodes of electric propulsion systems as the electron source. Hollow cathode technology is mature, with high emission current and low electron energy, but it requires the consumption of inert gas working fluids such as xenon, and necessitates the configuration of complex storage tanks, valves, and flow control systems. Its disadvantages include high power consumption (hundreds of watts), heavy weight (tens of kilograms), and system complexity.

[0010] Therefore, there is an urgent need for a low-power, small-size, propellant-free, lightweight, and simple-structured electronic current transmitting device to meet the functional, lifespan, and reliability requirements of space tethered propulsion systems.

[0011] Patent document CN109050994A discloses a propellantless electron-emitting spacecraft surface potential active controller, including a photocathode, a light source, an outer cylinder, and leads. The outer cylinder is open at both ends, with a transmissive photocathode at one end and a centrally open lead at the other. Both the photocathode and the lead are insulated from the outer cylinder. The light source is located outside the outer cylinder, and the photocathode generates photoelectrons under illumination. Positive and negative voltages are applied to the lead and the outer cylinder, respectively. This patent document uses a photocathode, and its basic principle is the absorption of external light source energy by a specific photoelectric material, exciting the electron emission process. This patent document is applicable to emission currents from nanoamps to microamps (10⁻⁶ Ω·cm). -9 ~10 -6A) Applications of phototubes and photomultiplier tubes with ranged electron currents belong to another field of cathode electronics. Reference: Lin Zulun, Wang Xiaoju, "Cathode Electronics", National Defense Industry Press, Beijing, January 2013, First Edition, P154. If the photocathode in this patent document is used in a space tethered propulsion system, assuming the current is microamps, for similar configurations and missions, the thrust is on the order of ~μN (one-thousandth), and the deorbit time is about 1000 years (thousand times). If the photocathode provides a milliamp (mA) or even larger electron current, it is necessary to increase the emission size of the photocathode to thousands of times, or to make an array of thousands of photocathodes to meet the mission requirements. At the same time, a giant specific light source is needed to meet the illumination requirements, which will greatly increase the difficulty of the manufacturing process and make engineering implementation difficult. In summary, the principle of the patent document with publication number CN109050994A is based on the photocathode and is not applicable to the field of electron current sources that require the range of milliamps to amperes. Summary of the Invention

[0012] To address the shortcomings of existing technologies, the purpose of this invention is to provide a propellant-free miniature thermionic emission device.

[0013] According to the present invention, a working fluid-free miniature thermionic emission device includes: a lead-out electrode, an electron emitter, a heating assembly, a shielding cover, an insulating connector, and a support cylinder;

[0014] The lead electrode is disposed outside the support cylinder, and the electron emitter and the heating assembly are disposed inside the support cylinder; the heating assembly is used to heat the electron emitter, so that the electron emitter emits thermionic current; the central axis of the electron emitter is disposed corresponding to the lead hole on the lead electrode, and there is a gap between the emitting end face of the electron emitter and the lead electrode;

[0015] The two ends of the insulating connector are respectively connected to the outer side wall of the support cylinder and the inner side wall of the lead electrode. The shielding cover is disposed on the periphery of the support cylinder, and there is a gap between the shielding cover and the support cylinder.

[0016] Preferably, one open end of the support cylinder is located inside the lead-out electrode, and the other open end of the support cylinder is located outside the lead-out electrode;

[0017] The electron emitter is disposed at the open end of the support cylinder located inside the lead electrode, and the power connection end of the heating assembly extends out of the open end of the support cylinder located outside the lead electrode.

[0018] Preferably, the electron emitter is an electron emitting material with low work function;

[0019] The electron emitter can be any one of the following: an oxide emitter, a barium tungsten emitter, or a lanthanum boride emitter.

[0020] Preferably, the emitting surface of the electron emitter is a concave spherical cap, a circular plane, or a polygonal plane.

[0021] Preferably, it also includes a power source and a heating power source;

[0022] The positive terminal of the power supply is connected to the lead-out terminal; the negative terminal of the power supply is connected to the negative terminal of the heating power supply and the negative terminal of the heating component; the positive terminal of the heating power supply is connected to the positive terminal of the heating component.

[0023] Preferably, the heating component is a spiral heater, and its configuration is any one of the following: planar single spiral, planar double spiral, three-dimensional single spiral, three-dimensional double spiral, or variable pitch three-dimensional single spiral structure.

[0024] Preferably, the insulating connector is a ceramic insulator.

[0025] Preferably, the heater is made of any one of the following materials: molybdenum, tantalum, niobium, tungsten, rhenium, or an alloy;

[0026] The alloy contains any two or more of the following materials: molybdenum, tantalum, niobium, tungsten, and rhenium.

[0027] Preferably, the support cylinder is made of any one of the following materials: molybdenum, tantalum, niobium, tungsten, rhenium, or alloys;

[0028] The alloy contains any two or more of the following materials: molybdenum, tantalum, niobium, tungsten, and rhenium.

[0029] Preferably, the shielding cover is made of any of the following materials: tantalum, niobium, tantalum-niobium alloy, molybdenum-rhenium alloy, foil, or strip.

[0030] Compared with the prior art, the present invention has the following beneficial effects:

[0031] 1. The thermionic emission device of the present invention is small in size, light in weight, low in power consumption, and can emit electron current in the range of hundreds of milliamperes. As an electron source for tethered propulsion systems, it does not require the consumption of propellant working fluid, can greatly improve the performance of propulsion systems, simplify system structure, reduce weight, and complete space missions such as orbital transfer or deorbiting at the end of the life of spacecraft.

[0032] 2. This invention can be used in space tethered propulsion systems to actively emit electron currents of tens to hundreds of milliamperes into the space environment. By collecting electrons in the space plasma environment through electron collection collisions, a current of a certain magnitude and direction can be formed in the conductive tether, forming a pathway with the charged ion layer in the space environment. The conductive tether generates a Lorentz force when cutting the Earth's magnetic field. This Lorentz force can provide thrust to the connected spacecraft. When the Lorentz force is in the same direction as the system's orbital velocity, it can provide thrust to the spacecraft without consuming any fuel, achieving orbital elevation. When the direction of the current in the tether is changed, the induced Lorentz force is opposite to the direction of the spacecraft's orbital velocity, thus achieving the drag function of space tethered propulsion, thereby lowering the spacecraft's orbit. Therefore, by adjusting the magnitude and direction of the current flowing into the tether, the magnitude and direction of the thrust generated by space tethered propulsion can be changed within a certain range.

[0033] 3. The device of the present invention does not require the consumption of working fluid, has a simple structure, small size, light weight, and emits low energy electronic current. The electronic current can be actively adjusted by heating power and lead-out voltage, which will bring huge economic benefits to the rope propulsion system.

[0034] 4. The propellantless electron source of the present invention can be used as a neutralizer for low-power electric thrusters (ion, Hall, field emission thrusters, etc.) to neutralize the plume of the thruster and prevent the spacecraft from becoming charged.

[0035] 5. The propellantless electronic source of the present invention can be used as a plasma contactor for active control of the surface potential of spacecraft such as space stations during space operations, to prevent the accumulation of local charges on the surface of spacecraft from causing electrostatic discharge, and to prevent electrostatic discharge from causing fires, ablation and other hazards to the safety of spacecraft surface materials and key equipment. Attached Figure Description

[0036] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0037] Figure 1 This is a schematic diagram of the structure of the working fluid-free micro thermionic emission device of the present invention;

[0038] Figure 2 This is a physical schematic diagram of the space rope propulsion system used in this invention.

[0039] Figure 3 This is a schematic diagram of the end face shape of the electron emitter of the present invention;

[0040] Figure 4 This is a configuration diagram of the heater of the present invention;

[0041] Figure 5 This is the circuit schematic diagram of the present invention.

[0042] The diagram shows:

[0043] Lead-out electrode 1, shield 4

[0044] Lead-out power supply 101 Insulating connector 5

[0045] Heating power supply 102, support cylinder 6

[0046] Electron emitter 2, lead-out hole 7

[0047] Heating component 3 Detailed Implementation

[0048] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0049] Example 1:

[0050] like Figures 1-5 As shown, this embodiment provides a propellant-free miniature thermionic emission device, including: a lead-out electrode 1, an electron emitter 2, a heating assembly 3, a shield 4, an insulating connector 5, and a support cylinder 6. The lead-out electrode 1 is disposed outside the support cylinder 6, and the electron emitter 2 and the heating assembly 3 are disposed inside the support cylinder 6. The heating assembly 3 is used to heat the electron emitter 2, causing the electron emitter 2 to emit thermionic current. The central axis of the electron emitter 2 is disposed corresponding to the lead-out hole 7 on the lead-out electrode 1. There is a gap between the emitting end face of the electron emitter 2 and the lead-out electrode 1. The two ends of the insulating connector 5 are respectively connected to the outer side wall of the support cylinder 6 and the inner side wall of the lead-out electrode 1. The shield 4 is disposed on the periphery of the support cylinder 6, and there is a gap between the shield 4 and the support cylinder 6.

[0051] The electron emitter 2, heating assembly 3, and support cylinder 6 are welded together as a single structure via flanges. The insulating connector 5 is a ceramic insulator.

[0052] The lead-out hole 7 is positioned directly opposite the central axis of the electron emitter 2. There is a gap between the emitting end face of the electron emitter 2 and the lead-out electrode 1.

[0053] The electron emitter 2 is made of an electron-emitting material with a low work function. The electron emitter 2 can be any of the following: an oxide emitter, a barium tungsten emitter, or a lanthanum boride emitter. The emission surface of the electron emitter 2 is a plane or a concave spherical cap. In this embodiment, the emission surface of the electron emitter 2 is a concave spherical cap, a circular plane, or a polygonal plane.

[0054] The working fluidless micro thermionic emission device of this embodiment also includes a lead-out power supply 101 and a heating power supply 102. The positive terminal of the lead-out power supply 101 is connected to the lead-out terminal 1, the negative terminal of the lead-out power supply 101 is connected to the negative terminal of the heating power supply 102 and the negative terminal of the heating component 3 respectively, and is grounded. The positive terminal of the heating power supply 102 is connected to the positive terminal of the heating component 3.

[0055] One open end of the support cylinder 6 is located inside the lead electrode 1, and the other open end of the support cylinder 6 is located inside the lead electrode 1. The electron emitter 2 is disposed at the open end of the support cylinder 6 located inside the lead electrode 1, and the power connection end of the heating component 3 extends out of the open end of the support cylinder 6 located outside the lead electrode 1.

[0056] Heating component 3 is a spiral heater. The heater is made of any one of the following materials: molybdenum, tantalum, niobium, tungsten, rhenium, or an alloy, wherein the alloy contains any two or more of the following materials: molybdenum, tantalum, niobium, tungsten, or rhenium. The heater configuration is any one of the following: planar single spiral, planar double spiral, three-dimensional single spiral, three-dimensional double spiral, or variable pitch three-dimensional single spiral structure.

[0057] The support cylinder 6 is made of any one of the following materials: molybdenum, tantalum, niobium, tungsten, rhenium, or an alloy, wherein the alloy contains any two or more of the following materials: molybdenum, tantalum, niobium, tungsten, or rhenium. The shielding cover 4 is made of any one of the following materials: tantalum, niobium, tantalum-niobium alloy, or molybdenum-rhenium alloy.

[0058] Example 2:

[0059] Those skilled in the art can understand this embodiment as a more specific description of Embodiment 1.

[0060] The technical solution of this embodiment relates to the field of aerospace propulsion technology. Specifically, it relates to the structural design of an electronic source for a spacecraft. In particular, it is applicable to tethered propulsion systems and can form a low-power propulsion system that does not require the consumption of propellant working fluid and can complete tasks such as spacecraft orbit transfer and deorbiting at the end of its life.

[0061] This embodiment provides a propellantless miniature thermionic emission device for a rope-driven propulsion system. The propellantless miniature thermionic emission device includes a lead-out electrode, an electron emitter, a heater, a shield, a ceramic insulator, and a support cylinder.

[0062] The heater is powered by an external power source to heat up, and the heat is conducted to the electron emitter, reaching the operating temperature to emit thermionic current. The lead-out electrode is electrically insulated from the electron emitter by a ceramic insulator. An appropriate lead-out voltage is applied to the lead-out electrode to form an electric field, which is beneficial for electron extraction. Through miniaturization design and material optimization, the aforementioned fluidless micro thermionic emission device can obtain tens to hundreds of milliamperes of electron current when supplied with a heating power of 5-20W.

[0063] The electron emitter is mounted at one end of a support cylinder. Inside the support cylinder, next to the emitter, a spiral heater is installed to facilitate heat transfer from the heater to the electron emitter via conduction. An external power source powers the heater, causing it to heat up. This heat is then transferred to the electron emitter via conduction until it reaches its operating temperature, emitting thermionic current. The support cylinder is surrounded by a shield with a gap between it and the support cylinder to reduce heat radiation loss and power consumption. The lead-out electrode is electrically insulated from the electron emitter by a ceramic insulator. A small hole on the lead-out electrode faces the central axis of the emitter, facilitating electron extraction. In this propellantless miniature thermionic emission device, the electron emitter, heater, support cylinder, and flange are welded together as a single structure, which improves structural strength and resistance to environmental forces.

[0064] The corded propulsion system uses a propellantless micro thermionic emission device, which is a thermionic source operating in a vacuum environment with a vacuum level better than 2×10⁻⁶. -2 Pa.

[0065] The corded propulsion system uses a propellantless micro thermionic emission device. The emitter is made of an electron-emitting material with a low work function. The emitter material can be one of oxide emitters, barium tungsten emitters, or lanthanum boride emitters, or other materials with low work function. The emission current density j0 [A / cm²] 2 This can be described by the Charleson formula:

[0066]

[0067] Where A0 is the theoretical value of the emission constant, which is taken as 120.4 A·cm for barium tungsten emitters. -2 ·K -2 T is the operating temperature of the emitter [K], φ c Let be the surface work function of the cathode emitter [eV], which is 2.12eV for a barium-tungsten emitter (1eV = 1.6 × 10⁻¹⁹ J), and k be the Boltzmann constant 1.38 × 10⁻¹⁹ J. -23 J / K, when the barium tungsten emitter operates at temperatures between 950℃ and 1100℃, the emission current density is calculated using the Charleson formula as follows:

[0068] j0 = 0.33~3.8A / cm 2

[0069] The emitting surface of an electron emitter can be a concave spherical cap, a circular plane, or a polygonal plane. A concave surface is advantageous for electron focusing and emission.

[0070] The heater is made of high-temperature resistant molybdenum, tantalum, niobium, tungsten, rhenium, or alloys thereof.

[0071] The heater configuration is any of the following: planar single helix, planar double helix, three-dimensional single helix, three-dimensional double helix, or variable pitch three-dimensional single helix structure.

[0072] The support cylinder is made of high-temperature resistant molybdenum, tantalum, niobium, tungsten, rhenium, or alloys of the above materials.

[0073] The heat shield is made of high-temperature resistant materials, preferably tantalum, niobium, tantalum-niobium alloy, molybdenum-rhenium alloy foil or strip.

[0074] A certain gap is maintained between the lead-out electrode and the emitting end face of the electron emitter, and a positive voltage is applied to draw out electrons.

[0075] Example 3:

[0076] Those skilled in the art can understand this embodiment as a more specific description of Embodiment 1.

[0077] This invention provides a propellant-free miniature thermionic emission device for a tethered propulsion system, the structure of which is as follows: Figure 1 As shown, it includes a lead-out electrode 1, an electron emitter 2, a heating assembly 3, a shielding cover 4, an insulating connector 5, and a support cylinder 6. The heating assembly 3 is a heater, and the insulating connector 5 is a ceramic insulator.

[0078] A planar electron emitter 2 is mounted at one end of a support cylinder 6. A spiral heating element 3 is installed inside the support cylinder 6 next to the emitter, facilitating heat transfer from the heating element 3 to the electron emitter 2 via conduction. A shielding cover 4 is located outside the support cylinder 6, with a gap between the shielding cover 4 and the support cylinder 6 to reduce heat radiation loss and power consumption. The lead-out electrode 1 is electrically insulated from the electron emitter 2 via an insulating connector 5. The lead-out electrode 1 has a small hole aligned with the central axis of the emitter, facilitating electron extraction. In this propellantless micro thermionic emission device, the electron emitter, heater, and support cylinder are welded together as a single structure, which improves structural strength and increases resistance to mechanical environments.

[0079] The emitter uses an electron-emitting material with a low work function, and the emission current density can be described by the Charleson formula. The emitter material is a barium-tungsten emitter, and the emission current density j0 [A / cm²]. 2 It can be calculated using equation (2):

[0080]

[0081] Where A0 is the theoretical value of the emission constant, which is taken as 120.4 A·cm for barium tungsten emitters. -2 ·K -2 T is the operating temperature of the emitter [K], φ cLet be the surface work function of the cathode emitter [eV], which is 2.12eV for a barium-tungsten emitter (1eV = 1.6 × 10⁻¹⁹ J), and k be the Boltzmann constant 1.38 × 10⁻¹⁹ J. -23 J / K, when the barium tungsten emitter operates at temperatures between 950℃ and 1100℃, the emission current density is calculated using the Charleson formula as follows:

[0082] j0 = 0.33~3.8A / cm 2

[0083] The electron emitter is a circular plane with a diameter of 5 mm, such as Figure 3 As shown in (2).

[0084] The heater heating wire is made of tungsten-rhenium wire with a diameter of 0.2mm to 0.4mm, and its configuration is one of the following: planar single helix, planar double helix, three-dimensional single helix, three-dimensional double helix, or variable pitch three-dimensional single helix structure. Figure 4 As shown.

[0085] The support cylinder is made of molybdenum, and the shielding cover is made of tantalum foil or molybdenum-rhenium foil.

[0086] A gap of 1-2 mm is maintained between the lead-out electrode and the emitting end face of the electron emitter, and a positive voltage is applied to draw out electrons.

[0087] The schematic diagram of the working circuit of the propellantless miniature thermionic emission device is shown below. Figure 5 The working environment vacuum level is not less than 2×10 -2 Pa. When the heating power is 5-20W and the heating time is about 5 minutes, the temperature of the electron emitter reaches 950℃-1150℃. When a voltage of 100-500V is applied to the lead-out electrode, an electron current can be observed to be generated. When the emitted electron current can reach 5-100mA, it can meet the electron source requirements of the rope propulsion system.

[0088] The electronic emission device of the present invention is small in size, light in weight, low in power consumption, and can emit electronic currents in the range of hundreds of milliamperes. As an electronic source for tethered propulsion systems, it does not require the consumption of propellant working fluid, can significantly improve the performance of propulsion systems, simplify system structure, reduce weight, and complete space missions such as orbital transfer or deorbiting at the end of the spacecraft's life.

[0089] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0090] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A propellant-free miniature thermionic emission device, characterized in that, include: Lead-out electrode (1), electron emitter (2), heating assembly (3), shielding cover (4), insulating connector (5), support cylinder (6); The lead-out electrode (1) is disposed outside the support cylinder (6), and the electron emitter (2) and the heating assembly (3) are disposed inside the support cylinder (6); the heating assembly (3) is used to heat the electron emitter (2) so that the electron emitter (2) emits thermionic current; the central axis of the electron emitter (2) is disposed corresponding to the lead-out hole (7) on the lead-out electrode (1), and there is a gap between the emitting end face of the electron emitter (2) and the lead-out electrode (1); The two ends of the insulating connector (5) are respectively connected to the outer side wall of the support cylinder (6) and the inner side wall of the lead electrode (1). The shield (4) is disposed on the periphery of the support cylinder (6), and there is a gap between the shield (4) and the support cylinder (6). The electron emitter (2), the heating component (3), and the support cylinder (6) are welded together to form a single structure. One open end of the support cylinder (6) is located inside the lead-out electrode (1), and the other open end of the support cylinder (6) is located outside the lead-out electrode (1); The electron emitter (2) is disposed at the open end of the support cylinder (6) located inside the lead-out electrode (1), and the power connection end of the heating component (3) extends out of the open end of the support cylinder (6) located outside the lead-out electrode (1).

2. The propellantless micro thermionic emission device according to claim 1, characterized in that, The electron emitter (2) is made of an electron emission material with a low work function. The electron emitter (2) can be any one of the following: oxide emitter, barium tungsten emitter, or lanthanum boride emitter.

3. The propellantless micro thermionic emission device according to claim 1, characterized in that, The emitting surface of the electron emitter (2) is a concave spherical cap, a circular plane, or a polygonal plane.

4. The propellantless micro thermionic emission device according to claim 1, characterized in that, It also includes a power supply (101) and a heating power supply (102). The positive terminal of the lead-out power supply (101) is connected to the lead-out terminal (1); the negative terminal of the lead-out power supply (101) is connected to the negative terminal of the heating power supply (102) and the negative terminal of the heating component (3); the positive terminal of the heating power supply (102) is connected to the positive terminal of the heating component (3).

5. The propellantless micro thermionic emission device according to claim 1, characterized in that, The heating component (3) is a spiral heater, and its configuration is any one of the following: planar single spiral, planar double spiral, three-dimensional single spiral, three-dimensional double spiral, or variable pitch three-dimensional single spiral structure.

6. The propellantless micro thermionic emission device according to claim 1, characterized in that, The insulating connector (5) is a ceramic insulator.

7. The propellantless micro thermionic emission device according to claim 5, characterized in that, The heater is made of any one of the following materials: molybdenum, tantalum, niobium, tungsten, rhenium, or alloys; The alloy contains any two or more of the following materials: molybdenum, tantalum, niobium, tungsten, and rhenium.

8. The propellantless micro thermionic emission device according to claim 1, characterized in that, The support cylinder (6) is made of any of the following materials: molybdenum, tantalum, niobium, tungsten, rhenium, or alloy; The alloy contains any two or more of the following materials: molybdenum, tantalum, niobium, tungsten, and rhenium.

9. The propellantless micro thermionic emission device according to claim 1, characterized in that, The shield (4) is made of any of the following materials: tantalum, niobium, tantalum-niobium alloy, molybdenum-rhenium alloy.

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