CN119429186B - A microwave detonation dual-mode space thruster - Google Patents
A microwave detonation dual-mode space thrusterInfo
- Publication number
- CN119429186B CN119429186B CN202411733150.3A CN202411733150A CN119429186B CN 119429186 B CN119429186 B CN 119429186B CN 202411733150 A CN202411733150 A CN 202411733150A CN 119429186 B CN119429186 B CN 119429186B
- Authority
- CN
- China
- Prior art keywords
- microwave
- detonation
- propulsion module
- thrust
- antenna
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/40—Arrangements or adaptations of propulsion systems
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/40—Arrangements or adaptations of propulsion systems
- B64G1/402—Propellant tanks; Feeding propellants
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/40—Arrangements or adaptations of propulsion systems
- B64G1/411—Electric propulsion
- B64G1/413—Ion or plasma engines
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Remote Sensing (AREA)
- Aviation & Aerospace Engineering (AREA)
- Physics & Mathematics (AREA)
- Plasma & Fusion (AREA)
- Plasma Technology (AREA)
Abstract
The invention discloses a microwave detonation dual-mode space thruster, which comprises a propellant supply system, a high-thrust propulsion module and a micro-thrust propulsion module which are coaxially arranged in sequence from inside to outside, wherein the high-thrust propulsion module comprises an injector, a plasma ignition combustion-supporting device, a detonation chamber, a tail nozzle and the like. The micro-thrust propulsion module comprises a microwave supply system, a magnetic pole, an air collecting cavity, a discharge chamber and a grid electrode, wherein the microwave supply system comprises a microwave connector and an antenna, the antenna is a cylindrical antenna, a convolution cavity is formed between the antenna and the magnetic pole, and outer electrons in electric working molecules in the convolution cavity do convolution motion in the convolution cavity, and resonance occurs when the convolution angular frequency is equal to the microwave frequency. The invention integrates the microwave discharge technology, the plasma combustion-supporting technology and the pulse detonation combustion technology, realizes the efficient and convenient switching of the working mode through circuit regulation, ensures that the thruster can meet the severe requirements of high thrust, high specific impulse, long service life and the like, and can be applied to diversified space tasks.
Description
Technical Field
The invention relates to the technical field of space propulsion of spacecrafts, in particular to a microwave detonation dual-mode space thruster.
Background
Along with the rapid development of aerospace industry in China, diversified space tasks put forward higher and more complex requirements on a propulsion system. The current space thruster is mainly divided into two types, namely chemical high-thrust propulsion and ionization high-specific impulse propulsion, and is difficult to meet the needs of diversified space tasks such as high thrust, high specific impulse and the like. The microwave ion electric propulsion technology is characterized in that microwave energy is used for breaking down gas to generate plasma, an electrostatic field is used for accelerating ions again to generate thrust, the microwave ion electric propulsion technology has the characteristics of wide thrust range, strong controllability, high specific impulse and capability of taking various gases as working media, has remarkable advantages in the aspects of service life and accurate thrust control, is limited by smaller thrust, is mainly applied to tasks such as orbit, gesture maintenance and deep space exploration of a spacecraft at present, and does not meet the requirements of rapid maneuvering of the spacecraft. Compared with the prior art, the pulse detonation chemical propulsion technology is a novel propulsion mode which is characterized by simple structure, low cost, high thrust weight ratio and high thermal cycle efficiency, is a novel propulsion mode which is characterized by high temperature and high pressure gas generated by pulse detonation waves and provides thrust, and for a spacecraft with large mass and rapid maneuvering, a pulse detonation engine is ideal choice, but is limited by chemical energy of the propellant and wall temperature, the pulse detonation engine has lower specific impact, and the requirement of continuously changing thrust cannot be realized. The plasma ignition combustion-supporting technology is a novel ignition combustion-supporting technology which utilizes the high temperature effect, chemical effect and pneumatic effect of plasma to improve the ignition capability and combustion efficiency, can strengthen ignition, improve the flame propagation speed and stability, widen the flameout limit, and further realize high-efficiency combustion under the conditions of higher air pressure, lower flame temperature and leaner oil.
The invention integrates the microwave discharge technology, the plasma combustion-supporting technology and the pulse detonation combustion technology, realizes the efficient and convenient switching of the working mode through circuit regulation, ensures that the thruster can meet the severe requirements of high thrust, high specific impulse, long service life and the like, and can be applied to diversified space tasks.
Disclosure of Invention
The invention aims to solve the technical problems of the prior art, and provides a microwave detonation dual-mode space thruster which integrates a microwave discharge technology, a plasma combustion supporting technology and a pulse detonation combustion technology, so that the problem of long pulse detonation distance is effectively solved, the size of the thruster is reduced, the ignition detonation performance is improved, and meanwhile, the efficient and convenient switching of the working mode is realized through circuit regulation and control, so that the thruster can meet severe requirements of high thrust, high specific impulse, long service life and the like, and can be applied to diversified space tasks.
In order to solve the technical problems, the invention adopts the following technical scheme:
a microwave detonation dual-mode space thruster comprises a propellant supply system, a high-thrust propulsion module and a micro-thrust propulsion module which are coaxially arranged in sequence from inside to outside.
The propellant supply system comprises a propellant storage tank, a detonation branch pipeline and an ionization branch pipeline, wherein the detonation branch pipeline and the ionization branch pipeline are respectively connected with the propellant storage tank, the detonation branch pipeline can supply fuel and oxidant to the high-thrust propulsion module, and the ionization branch pipeline can supply electric working media to the micro-thrust propulsion module.
The micro-thrust propulsion module comprises a microwave supply system, a magnetic pole, an air collecting cavity, a discharge chamber and a grid electrode.
The gas collecting cavity, the discharge chamber and the grid electrode are coaxially arranged in sequence from front to back.
The magnetic pole is a magnetic ring and is coaxially arranged at the head of the discharge chamber adjacent to the gas collecting cavity.
The microwave supply system includes a microwave connector and an antenna.
The antenna is a cylindrical antenna and is coaxially inserted on the inner wall of the discharge chamber positioned at the inner side of the magnetic pole, a convolution cavity is formed between the antenna and the magnetic pole, and the convolution cavity is respectively communicated with the gas collecting cavity and the discharge chamber.
The microwave connector is used for feeding microwaves to the antenna.
The electric medium provided by the ionization branch pipeline enters the convolution cavity after passing through the gas collecting cavity, at the moment, outer electrons in electric medium molecules do convolution motion in the convolution cavity, resonance occurs when the convolution angular frequency omega e is equal to the microwave frequency f, and the convolution angular frequency omega e is matched with the magnetic field intensity B of the magnetic pole.
The calculation formula of the convolution angle frequency omega e is as follows:
wherein e and m e are the charge and mass, respectively, of the outer electrons in the electrical property molecule.
The magnetic pole is an electromagnet, and the magnetic field intensity B is adjusted by changing the exciting current of the magnetic pole so as to be matched with the microwave frequency f, when the microwave frequency f is 2.45GHz, the magnetic field intensity B is adjusted to be 0.875T, and when the microwave frequency f is 4.2GHz, the magnetic field intensity B is adjusted to be 1.5T.
The outer shell of the high-thrust propulsion module is an inner shell of a discharge chamber in the micro-thrust propulsion module, and the outer shell of the high-thrust propulsion module is made of high-temperature resistant insulating materials.
The high-thrust propulsion module comprises an injector, a plasma ignition combustion-supporting device, a detonation chamber and a tail nozzle, wherein the injector, the detonation chamber and the tail nozzle are coaxially arranged in sequence from front to back, and the plasma ignition combustion-supporting device is arranged on the side wall of the front end of the detonation chamber.
A DDT enhancement device is arranged in the knocking chamber.
The knocking branch pipe is provided with a check valve, when the pressure in the knocking chamber is lower than the injection pressure, the check valve is opened, otherwise, the check valve is closed.
The axial length of the antenna is greater than that of the magnetic pole, and the antenna and the magnetic pole are flush at the tail end surface of the discharge chamber.
When the spacecraft needs to execute the space task of maintaining the orbit and adjusting the attitude, the micro-thrust propulsion module independently works to realize high specific impulse.
When the micropower propulsion module works independently, ions and electrons are alternately extracted by applying a periodically-changing voltage to the grid electrode, so that self-neutralization is realized, wherein the application frequency of the grid electrode voltage is larger than the minimum extraction frequency of the grid electrode.
The invention has the following beneficial effects:
1. the invention fully utilizes the advantages of a microwave ionization technology, a plasma combustion supporting technology and a detonation combustion technology, and designs the microwave detonation dual-mode space thruster. The thruster can work in two different modes of high thrust and micro thrust, and can combine the advantages of high specific thrust and high thrust of pulse detonation chemical propulsion technology of microwave ion electric propulsion technology, so that the thruster has wider performance coverage range and more flexible space task application.
2. The micro-thrust propulsion module and the large-thrust propulsion module can share the propellant, and the propellant is selected from fuel and oxidant such as methane, kerosene, ammonia, oxygen, nitrous oxide, hydrogen peroxide and the like, so that the mass-volume ratio of the storage tank and corresponding matched pipelines is reduced, and space resources such as mass, volume and the like can be saved for the spacecraft.
3. The invention adopts a modularized design, the micro-thrust propulsion module and the large-thrust propulsion module can work independently, and the high temperature of the wall surface can provide favorable conditions for starting the other working mode when the working mode is switched. The coaxial nested space structure of the thruster is convenient for replacing and recombining other types of propulsion units according to requirements when being applied to different scenes in the future.
4. The invention carries out structural optimization design on the micro-thrust propulsion module, provides a coaxial annular discharge chamber configuration, adopts a cylindrical antenna and a circular grid system, has larger coupling area ratio of microwaves and electromagnetic fields and higher coupling degree compared with the cylindrical configuration design, can effectively enhance ionization and acceleration extraction processes of propellant molecules, and realizes the performance optimization of quick start, high specific impulse, high adjustment precision and the like.
5. According to the invention, through the periodical change of the grid voltage, ions and electrons are alternately led out, the self-neutralization effect is realized, an external neutralizer is not needed, the structure of the thruster is simplified, the restriction of particles on the pollution of the traditional neutralizer hot cathode material can be avoided, the failure rate is reduced, the service life of the thruster is prolonged, and the emission cost is reduced.
6. For the high-thrust propulsion module, the invention adopts a transition (DDT) mode of slow combustion to detonation to indirectly detonate, and is provided with a DDT reinforcing device to accelerate the formation of detonation waves, and adopts a one-way valve to realize the self-adaptive control of the pulse detonation circulation process, so that an additional control device is not required to be additionally arranged, and the complexity of the system is reduced.
7. The invention applies the plasma combustion-supporting technology to pulse detonation chemical propulsion, adopts a plasma ignition combustion-supporting device, can adopt a DBD plasma ignition combustion-supporting device, an arc plasma ignition combustion-supporting device, a microwave plasma ignition combustion-supporting device and the like, preferably adopts a microwave plasma ignition combustion-supporting device, can share a microwave excitation source and a power supply, has a simplified structure, promotes combustion initiation, and realizes mutual coupling of ignition and combustion supporting. Meanwhile, according to practical application requirements, the invention provides 2n (n=1, 2) ignition pairs, so that uniform ignition is realized, a large amount of active particles are generated, and stable combustion is promoted.
Drawings
Fig. 1 shows a three-dimensional perspective view of a microwave detonation dual-mode spatial thruster of the present invention.
Fig. 2 shows a three-dimensional cross-sectional view of a microwave detonation dual-mode spatial thruster of the present invention.
Fig. 3 shows an explosion schematic of a microwave detonation dual mode spatial thruster of the present invention.
Fig. 4 shows a three-dimensional cross-section of a high thrust propulsion module according to the invention.
Fig. 5 shows a three-dimensional cross-section of a micro-thrust propulsion module according to the invention.
Fig. 6 shows a flow chart of the adaptive duty cycle of the high thrust propulsion module of the present invention.
The method comprises the following steps:
10. A propellant supply system;
11. propellant storage tank, ionization branch pipeline, knocking branch pipeline and one-way valve;
20. a micro-thrust propulsion module;
21. a microwave supply system 211, a microwave connector 212, an antenna;
22. Magnetic pole, 23, gas collecting cavity, 231, distributing hole, 24, discharge chamber, 25, grid, 26, shielding shell, 27, discharge chamber shell;
30. a high thrust propulsion module;
31. Injector, 32, plasma ignition combustion improver, 33, detonation chamber, 34, tail nozzle and 35, DDT reinforcing device.
Detailed Description
The invention will be described in further detail with reference to the accompanying drawings and specific preferred embodiments.
In the description of the present invention, it should be understood that the terms "left", "right", "upper", "lower", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, are merely for convenience in describing the present invention and simplifying the description, and do not indicate or imply that the apparatus or element being referred to must have a specific orientation, be configured and operated in a specific orientation, and "first", "second", etc. do not indicate the importance of the components, and thus are not to be construed as limiting the present invention. The specific dimensions adopted in the present embodiment are only for illustrating the technical solution, and do not limit the protection scope of the present invention.
As shown in fig. 1,2 and 3, a microwave detonation dual-mode space thruster comprises a propellant supply system 10, a high-thrust propulsion module 30 and a micro-thrust propulsion module 20 which are coaxially arranged in sequence from inside to outside.
As shown in fig. 2, the propellant supply system comprises a propellant tank 11, and detonation branch lines 13 and ionization branch lines 12 respectively connected to the propellant tank.
The propellant storage tank is shared by the high-thrust propulsion module 30 and the micro-thrust propulsion module 20, so that the space resources of the spacecraft are fully utilized, and meanwhile, the effective saving of the quality resources is realized. The propellant tanks preferably include a fuel tank and an oxidizer tank. The fuel tank preferably stores methane, coal oil, or the like, and the oxidizing agent tank preferably stores oxygen, or the like.
The detonation branch lines described above are capable of supplying fuel and oxidant to the high thrust propulsion module, and are preferably provided with a one-way valve 14.
The ionization branch pipeline can supply an electric working medium to the micro-thrust propulsion module. The electric working medium is preferably one or a combination of fuel and oxidant. Further, the ionization branch line is preferably provided with a pyrolysis gasification device.
As shown in fig. 2,3 and 5, the micro-thrust propulsion module, in a coaxial annular configuration, preferably includes a microwave supply system 21, a magnetic pole 22, an air collection chamber 23, a discharge chamber 24, a grid 25, a shield 26 and a discharge chamber housing 27.
The gas collecting cavity, the discharge chamber and the grid electrode are coaxially arranged in sequence from front to back.
The magnetic pole is a magnetic ring and is coaxially arranged at the head of the discharge chamber adjacent to the gas collecting cavity.
The microwave supply system includes a microwave connector and an antenna.
The antenna is a cylindrical antenna and is coaxially inserted on the inner wall of the discharge chamber positioned at the inner side of the magnetic pole, and a convolution cavity is formed between the antenna and the magnetic pole. Further, the length of the antenna extending into the discharge chamber is preferably in the range of 1/2-2/3 of the discharge chamber, slightly greater than the length of the magnetic poles. In this embodiment, the axial length of the antenna is greater than the axial length of the magnetic pole, and the antenna and the magnetic pole are flush at the tail end surface of the discharge chamber.
As shown in fig. 5, a circle of uniformly distributed distribution holes 231 are formed on the tail end surface of the gas collecting cavity, the gas collecting cavity is communicated with the convolution cavity through the distribution holes, and meanwhile, the convolution cavity is directly communicated with the discharge chamber. The retention time of the electric working medium entering the gas collection cavity is prolonged due to the influence of the blocking force, and the gas is uniformly distributed and introduced into the discharge chamber through the distribution holes.
The microwave connector is used for feeding microwaves to the antenna.
The magnetic pole is preferably an electromagnet, and the magnetic field intensity B is adjusted by changing the exciting current of the magnetic pole so as to be matched with the microwave frequency f, wherein the magnetic field intensity B is adjusted to be 0.875T when the microwave frequency f is 2.45GHz, and the magnetic field intensity B is adjusted to be 1.5T when the microwave frequency f is 4.2 GHz.
The outer periphery of the discharge cell casing 27 is preferably provided with a shield case 26.
As shown in fig. 4, the high thrust propulsion module, in a cylindrical configuration, mainly comprises an injector 31, a plasma ignition booster 32, a detonation chamber 33 and a tail pipe 34.
The injector, the detonation chamber and the tail nozzle are coaxially arranged from front to back in sequence, wherein the injector is preferably a coaxial injector, so that the mixing and detonation process of fuel and oxidant can be promoted.
The igniter is arranged on the front side wall of the detonation chamber. The igniter adopts a plasma ignition combustion-supporting device, and can adopt a DBD plasma ignition combustion-supporting device, an arc plasma ignition combustion-supporting device, a microwave plasma ignition combustion-supporting device and the like. According to practical application needs, the plasma ignition combustion-supporting devices are arranged in pairs for 2n (n=1, 2..) to realize uniform ignition and generate a large amount of active particles to promote stable combustion.
The detonation and knocking adopts a transition (DDT) indirect detonation mode from slow combustion to knocking. The DDT reinforcing device 35 is arranged in the detonation tube to accelerate the transition from slow combustion to detonation, and Shchelkin spiral wires, semicircular disc-shaped bulges, blocking plates, pore plates and the like can be adopted.
Further, the arrangement of the check valve can be used for realizing the self-adaptive control of the knocking circulation process as shown in fig. 6, and the opening and closing of the check valve are completely self-adaptively controlled by the working process of the knocking circulation, so that special control equipment like a solenoid valve is not needed. The check valve is opened when the pressure in the detonation chamber is lower than the injection pressure, and is closed otherwise.
The tail nozzle preferably employs a converging-diverging nozzle that maintains the detonation chamber at a sufficient pressure to accommodate high vacuum conditions.
In connection with the practice of the invention, the temperature of the propulsion module will rise sharply when in operation, and the wall material is to be optimized. The high-temperature resistant materials such as titanium alloy can be selected for the micro-thrust propulsion module shell, the insulating materials with good high-temperature resistance such as silicon-based ceramics and carbon-carbon composite materials can be selected for the high-thrust propulsion module shell, and the high-temperature resistant materials are suitable for the high-temperature environment of chemical high-thrust combustion and can meet the conditions for generating and maintaining the plasma of the outer ring ionization micro-thrust propulsion module. When one of the modules works, the wall surface temperature is higher, the working medium ionization or combustion process of the other module can be promoted, and favorable conditions are created for the switching and starting of the working modes of the thruster.
The invention can make the high-thrust propulsion module and the micro-thrust propulsion module work independently.
1. High thrust propulsion
When the spacecraft needs to execute space orbit transfer and other rapid maneuvering tasks, the high-thrust propulsion module works independently to realize high thrust, and the specific working cycle process is as follows:
(1) Propellant filling, namely opening a fuel and oxidant switch and a one-way valve, and filling the detonation tube according to a certain proportion.
(2) And (3) igniting, namely starting the plasma ignition combustion-supporting device, igniting according to a given frequency, and simultaneously promoting auxiliary combustion.
(3) The detonation wave is initiated, formed and propagated, the gas is burnt in a deflagration mode after being ignited, slow combustion wave is generated in the detonation tube firstly, after DDT, the slow combustion wave becomes a high-speed stable detonation wave, and then the detonation wave follows the expansion wave. The check valve automatically closes when the pressure in the detonation chamber is higher than the injection pressure.
(4) The detonation wave is transmitted out, namely the detonation wave propagates to the outlet end and is further expanded and discharged through the tail nozzle, and meanwhile, great thrust is generated.
(5) And the expansion wave is generated at the pipe orifice and is transmitted into the detonation chamber, the burnt gas is emptied, and the gas parameters in the detonation chamber are gradually balanced.
(6) The cycle repeats, with the check valve reopened when the pressure in the detonation chamber is below the injection pressure, and the above duty cycle repeats.
2. Thrust propulsion with micro-thrust
When the spacecraft needs to execute space tasks such as orbit maintenance, attitude adjustment and the like, the micro-thrust propulsion module works independently to realize high specific impulse, and the specific work cycle process is as follows:
(1) Magnetic field and microwave feeding, namely, starting the magnetic pole and the microwave source to generate a magnetic field and feeding microwaves into the discharge chamber through the antenna.
(2) And (3) propellant is introduced, namely a propellant storage and supply system is started, and propellant is introduced.
(3) And electron cyclotron resonance heating, namely under the combined action of microwaves and magnets, outer electrons of propellant molecules in the discharge chamber do cyclotron motion, and the cyclotron angular frequency is as follows:
Wherein e and m e are the charge and mass of electrons, respectively, and B is the magnetic field strength.
When the frequency of omega e is equal to that of externally applied microwaves, electrons and microwaves resonate, microwave energy is continuously coupled into the electrons, the electron temperature is increased, the electron energy level is increased, the electrons are converted into high-energy electrons, and finally the electrons fly out of the constraint of atomic nuclei to realize ionization.
(4) The plasma is generated by the collision of the generated high-energy electrons with the propellant molecules introduced into the discharge chamber to ionize the propellant molecules, so as to generate the plasma.
(5) And (3) accelerating and spraying the generated plasma by a grid system to generate thrust. At this time, ions and electrons are alternately extracted by applying a periodically varying voltage to the gate electrode, wherein the applied frequency of the gate voltage is greater than the minimum extraction frequency of the gate electrode.
The calculation formula of the minimum extraction frequency f min of the grid is preferably:
wherein:
Wherein L eff is the effective distance of the ion moving between the grids, L g is the grid spacing, t s is the screen grid thickness, r s is the grid hole radius, U s is the grid voltage, M is the mass of the ion, and e is the electron charge.
When the double-grid structure is adopted, the double-grid spacing is 0.5mm, the screen grid thickness is 0.4mm, the screen grid hole radius is 1.0mm, the minimum extraction frequency corresponding to the 1200V grid voltage is 15.6073MHz, and the self-neutralization extraction frequency can be met only by applying the grid voltage which is larger than the minimum extraction frequency.
The gate extracts ions when the gate voltage is greater than zero and extracts electrons when the gate voltage is equal to or less than zero. The total amount of ions and electrons led out by the grid electrode is equal in average in each period, and the beam current is quasi-neutral, so that the self-neutralization effect can be realized, an external neutralizer is not needed, the complexity of the system can be reduced, and the failure rate is reduced.
The invention combines the microwave discharge technology, the plasma combustion-supporting technology and the pulse detonation combustion technology, integrates the technical advantages, effectively solves the problem of long pulse detonation distance, improves the ignition capability and the combustion efficiency, and realizes the efficient and convenient switching of the working mode through circuit regulation. The microwave ion electric propulsion technology has the advantages of wide thrust range, strong controllability and high specific impulse, can use various gases as working media, has remarkable advantages in the aspects of service life and accurate thrust control, has simple structure and low cost, has high thrust weight ratio and high thermal cycle efficiency, and can enhance ignition, improve flame propagation speed and stability and realize high-efficiency combustion under the conditions of higher air pressure, lower flame temperature and leaner oil. The invention realizes the combination of the high-thrust propulsion and the micro-thrust propulsion module, and innovates and designs a part component, and the thruster can work in a microwave ion electric propulsion mode and a pulse detonation chemical propulsion mode respectively, has the characteristics of high thrust, high specific impulse, long service life and the like, overcomes the application limitation of a single propulsion technology, can be applied to diversified space tasks, and provides a new thought for the combination of other types of propulsion units in the future.
The preferred embodiments of the present invention have been described in detail above, but the present invention is not limited to the specific details of the above embodiments, and various equivalent changes can be made to the technical solution of the present invention within the scope of the technical concept of the present invention, and all the equivalent changes belong to the protection scope of the present invention.
Claims (10)
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202411733150.3A CN119429186B (en) | 2024-11-29 | 2024-11-29 | A microwave detonation dual-mode space thruster |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202411733150.3A CN119429186B (en) | 2024-11-29 | 2024-11-29 | A microwave detonation dual-mode space thruster |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN119429186A CN119429186A (en) | 2025-02-14 |
| CN119429186B true CN119429186B (en) | 2025-07-15 |
Family
ID=94510643
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202411733150.3A Active CN119429186B (en) | 2024-11-29 | 2024-11-29 | A microwave detonation dual-mode space thruster |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN119429186B (en) |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO1997033790A1 (en) * | 1996-03-15 | 1997-09-18 | Wong Alfred Y | High-altitude lighter-than-air stationary platforms including ion engines |
| CN113294264A (en) * | 2021-04-16 | 2021-08-24 | 中国人民解放军战略支援部队航天工程大学 | Double-component variable-thrust rotary detonation rocket engine based on pintle injector |
Family Cites Families (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN100549399C (en) * | 2006-09-20 | 2009-10-14 | 西北工业大学 | A kind of high-frequency pulse pinking engine and controlling method thereof |
| GB0904850D0 (en) * | 2009-03-23 | 2009-05-06 | Rolls Royce Plc | Magneto-plasma-dynamic generator and method of operating the generator |
| CN205858522U (en) * | 2016-08-04 | 2017-01-04 | 广西玉柴机器股份有限公司 | The plasma system of natural gas engine |
| DE102022000797A1 (en) * | 2021-03-10 | 2022-09-15 | Mathias Herrmann | Ignition concept and combustion concept for engines and rockets; the most effective or directed excitation and ignition possible by means of adapted electromagnetic radiation or electromagnetic waves (e.g. radio waves, microwaves, magnetic waves) and catalytic absorbers to increase the energetic efficiency and thrust |
| CN115142983B (en) * | 2022-06-16 | 2024-08-27 | 中国人民解放军战略支援部队航天工程大学 | Spacecraft hybrid power thruster based on chemical-electric depth fusion |
-
2024
- 2024-11-29 CN CN202411733150.3A patent/CN119429186B/en active Active
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO1997033790A1 (en) * | 1996-03-15 | 1997-09-18 | Wong Alfred Y | High-altitude lighter-than-air stationary platforms including ion engines |
| CN113294264A (en) * | 2021-04-16 | 2021-08-24 | 中国人民解放军战略支援部队航天工程大学 | Double-component variable-thrust rotary detonation rocket engine based on pintle injector |
Also Published As
| Publication number | Publication date |
|---|---|
| CN119429186A (en) | 2025-02-14 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US4577461A (en) | Spacecraft optimized arc rocket | |
| US4548033A (en) | Spacecraft optimized arc rocket | |
| CN108869220B (en) | A propulsion device and propulsion method for a space maneuvering platform | |
| US7452513B2 (en) | Triple helical flow vortex reactor | |
| CN116771548B (en) | Chemical-electric arc combined power thruster for space and application method thereof | |
| CN115142983B (en) | Spacecraft hybrid power thruster based on chemical-electric depth fusion | |
| US7395656B2 (en) | Dual mode hybrid electric thruster | |
| US12183556B2 (en) | Methods and systems for increasing energy output in Z-pinch plasma confinement system | |
| CN110425045B (en) | continuous rotation detonation engine | |
| CN107061210B (en) | A kind of pulsed plasma thruster accelerated based on electrothermal and electromagnetic mixing | |
| CN114526499B (en) | A two-phase pulse detonation combustor based on rotating sliding arc ignition | |
| CN114810427B (en) | High-energy green liquid propeller ignition device and method for microwave excited plasma | |
| CN210106081U (en) | Solid ablation type magnetic plasma thruster | |
| RU2633075C1 (en) | Method for creating electric propulsion thrust | |
| US12035454B2 (en) | Plasma engine using ion extraction | |
| CN119429186B (en) | A microwave detonation dual-mode space thruster | |
| RU93962U1 (en) | Anti-aircraft guided missile | |
| CN110131120A (en) | Solid ablation type magnetic plasma thruster | |
| CN220434898U (en) | Rotary detonation engine device and aircraft | |
| JPH0771361A (en) | Production device for space navigation craft | |
| RU2675732C2 (en) | Hydrocarbon fuel combustion method and device for its implementation | |
| CN104314692A (en) | A Microwave Surface Wave Ignition and Combustion Supporting Device | |
| Martinez-Val et al. | Jet-ignited indirect-drive inertial fusion targets | |
| RU2374481C1 (en) | Liquid-propellant rocket engine with additional electromagnetic acceleration of working body | |
| RU2303156C1 (en) | Liquid-propellant rocket engine with additional electromagnetic acceleration of propulsive mass |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |