36 results about "Specific impulse" patented technology

The application discloses a kind of plasma propulsion devices, belong to plasma electric propulsion technical field, comprising: discharge chamber, spiral antenna I, center arc structure, spiral antenna II, center compensation structure, radio frequency power supply and electromagnetic coil;Center arc structure is located in discharge chamber, spiral antenna I spiral is wound in the outer wall of discharge chamber, with the position opposite of center arc structure;Center compensation structure is located in discharge chamber, spiral antenna II spiral is wound in the outer wall of discharge chamber, with the position opposite of center compensation structure;Spiral antenna I and spiral antenna II are electrically connected with radio frequency power supply respectively;Two electromagnetic coils are wound and set in the outer wall of discharge chamber, respectively located in the two ends of spiral antenna II;Center compensation structure can release neutral particles under the action of heat;The application can improve ion number density, realize high specific impulse output of propulsion device, improve propulsion efficiency.

The invention discloses an ultra-low orbit satellite electric propeller two-stage cascade power overall planning control method, and relates to the technical field of satellite electric propellers, and the method comprises the steps: constructing a thrust-radio frequency power matching model; utilizing an extended Kalman filtering algorithm to carry out radio frequency source power dynamic estimation technology research, and constructing a state observer; based on a thrust-radio frequency power matching model and a state observer, constructing a self-adaptive predictive power control framework based on measured data and dynamic state estimation; determining a power coupling mechanism between the spiral wave radio frequency source and the ion cyclotron resonance heating radio frequency source, and constructing a front-stage and back-stage power linkage relation mathematical model; based on a front-stage and back-stage power linkage mathematical model, a dynamic power optimal distribution method with optimal efficiency, thrust and specific impulse under power constraint is established, and a feedback controller based on delay compensation linear quadratic Gaussian control is constructed. And the performance of the ultra-low orbit satellite electric propeller is optimized.

This invention belongs to the field of engines, specifically relating to a fuel supply structure and a solid-liquid hybrid rocket engine, including an oxygen-enriched propellant supply mechanism and a gel fuel supply mechanism. The oxygen-enriched propellant supply mechanism includes a coating column, which includes a coating sleeve with an opening facing the combustion chamber, and the coating sleeve is used to hold the oxygen-enriched propellant. The gel fuel supply mechanism includes a gel fuel supply device, an annular channel, and an atomizer arranged sequentially. The annular channel is arranged around the outer edge of the coating sleeve, with its downstream end protruding downstream of the coating sleeve. The atomizer is located at the end of the annular channel protruding downstream of the coating sleeve, and the annular channel and the coating sleeve together seal the upstream end of the combustion chamber. This invention's engine adopts a novel combination of gel fuel and oxygen-enriched propellant, combining some of the advantages of both to achieve the effects of increased specific impulse, adjustable thrust, and multiple starts.

The present application relates to the technical field of aerospace engine, in particular to a liquid fuel engine thrust chamber needle bolt head regenerative cooling structure, comprising: a thrust chamber and a needle injection injector, the injector shell of the needle injection injector is arranged in the thrust chamber from the top of the thrust chamber, and the injector shell is located in the thrust chamber, at least one week main injection hole is arranged on the radial of the injector shell, the annular gap is formed between the injector shell and the top shell of the thrust chamber, and the end of the injector shell is provided with an inner recess structure in the thrust chamber; The first cooling flow channel and the second cooling flow channel are arranged in the injector shell above and below the main injection hole, and the outlet of the second cooling flow channel is connected with the rotational flow tangential hole; wherein, the liquid propellant in the propellant combination is injected into the injector shell, and the annular gap between the injector shell and the top shell of the thrust chamber is used for injecting another propellant in the propellant combination. The present application improves the combustion efficiency of the thrust chamber, and improves the specific impulse and the thrust-to-weight ratio of the engine.

The invention provides a space propulsion method of a nanosecond-femtosecond composite light filament, which is used for solving the problem of low long-distance laser propulsion efficiency. The method specifically comprises the following steps: forming a kilometer-level length filament through a space filament forming principle of femtosecond laser, stabilizing the power density at 1013-1015W / cm < 2 > magnitude by utilizing a light intensity clamping effect, and realizing dozens of kilometer-level low-loss energy transmission; meanwhile, the transmission direction of the nanosecond laser is adjusted through the optical lens group, the formed nanosecond filament and the femtosecond laser are coaxially transmitted or transmitted at a certain angle and act on the same position, a plasma channel generated by the filament enables the transmission loss of the high-energy nanosecond laser to be greatly reduced, and the energy absorption of the target material to the nanosecond laser is remarkably improved; therefore, the surface of the target material is melted, gasified and even ionized in an extremely short time, a large number of high-speed reverse-spraying plasma plumes are generated, and high specific impulse and thrust are obtained. The method is suitable for various target materials, and precise control over the thrust, the specific impulse and the propelling direction can be achieved by regulating and controlling laser parameters.

The present application relates to the technical field of spacecraft liquid rocket engine thrust chamber design, and provides a wide working condition liquid rocket engine thrust chamber, which comprises a combustion chamber and a direct impingement injector, the combustion chamber comprises a straight section, a converging section, a throat and an expansion section connected in sequence, the direct impingement injector is arranged at the end of the straight section, the length-diameter ratio of the straight section is 1.2-1.5, the characteristic length is 290-350 mm, the direct impingement injector is provided with a central zone first injection hole, a central zone second injection hole and a peripheral zone injection hole, the oxidizer is sprayed out through the central zone first injection hole, and the fuel is sprayed out through the central zone second injection hole and the peripheral zone injection hole.The present application greatly widens the working condition range that the thrust chamber can adapt to, and solves the problems that the throat temperature of a high specific impulse performance space engine is too high and is greatly affected by working condition deviation.

The invention relates to an efficient miniature ion thruster based on microwave ionization enhancement, and relates to the field of ion thrusters, a permanent magnet group comprises a plurality of annular permanent magnets which are distributed at intervals along the axis direction of a shell, and cusped magnetic fields are formed between adjacent permanent magnets to reduce the loss of plasmas on the wall surface; the tungsten filament cathode is fixedly connected to the middle of one side of the shell and the axis coincides with the axis of the shell; the annular anode is fixedly connected in the shell on the periphery of the tungsten filament cathode, the axis of the anode coincides with the axis of the tungsten filament cathode, and the gas supply assembly is fixedly connected to the shell on one side of the anode so as to supply a gaseous propellant into the shell; the device has the advantages that microwave energy coupling, efficient magnetic confinement of a cusped magnetic field, ionization area expansion and high specific impulse acceleration of a grid system are combined, and therefore efficient ionization and high specific impulse acceleration of low-density working medium gas can be achieved under the low-power condition; the device has the advantages of compact structure, high efficiency and high specific impulse.

The application discloses a kind of far-field uniform explosion load simulation loading device and test method, including explosion chamber, suspension piece, rectifier tube, target structure, pressure sensor, laser vibration meter and display terminal;Explosion chamber one end is open, the other end is connected rectifier tube, explosion chamber inner cavity and rectifier tube inner cavity are through;Suspension piece is set in explosion chamber top, and suspension piece can move back and forth along rectifier tube axis direction;Rectifier tube is used to shape explosion shock wave;Target structure is connected at the end of rectifier tube, pressure sensor is used to detect the pressure load of target structure surface, laser vibration meter is aimed at target structure, for measuring the structure response of target structure;The loading device of the application has the function of adjustable pressure peak value and specific impulse, can be used to carry out the damage effect research of target structure under the combined action of different pressure peak value, specific impulse, can provide support for the damage research of typical target structure under large equivalent charge far-field explosion load.

The application discloses a performance analysis method of a mixed fuel multi-working condition oblique detonation engine, and the oblique detonation engine comprises an air inlet channel section, a fuel mixing section, a combustion chamber section and a tail nozzle section, and the method comprises the following steps: according to the sequential characteristics of the working process of the oblique detonation engine, sequentially performing parameter solving of an oblique shock wave compression module of the air inlet channel section, parameter solving of a mixing module of the fuel mixing section, parameter solving of an oblique detonation module of the combustion chamber section and parameter solving of an isentropic expansion module of the tail nozzle section; and combining corresponding parameters of the solving processes, including temperature, pressure and velocity, performing theoretical thrust calculation of the oblique detonation engine, and calculating thrust performance including specific impulse. The method sets fuel composition, air inlet working condition, mixing working condition, combustion working condition and expansion working condition of the oblique detonation engine, and can quickly calculate specific impulse of the oblique detonation engine and each physical quantity along the way.

This utility model relates to the field of satellite propulsion equipment technology, and specifically provides an integrated gas-fuel hybrid propulsion system, including an integrated tank, gas generator, gas container, and thruster. The tank contains a flexible metal diaphragm, with one side forming a gas chamber and the other a liquid chamber. The gas chamber has a gas inlet, and the liquid chamber has a liquid inlet. The liquid inlet's output pipe is connected to the gas generator, the gas generator's output pipe is connected to the gas container, and the gas container's output pipe is connected to the thruster. This utility model solves the problems of low specific impulse in cold gas propulsion systems, heavy thruster mass and large envelope size in monocomponent chemical propulsion systems, and the difficulty in achieving low thrust output in bicomponent chemical propulsion systems, thus meeting the application requirements of satellites.

This invention relates to a multiphysics-coupled optimization design method and equipment for a toroidal pebble bed reactor core used for nuclear thermal propulsion. The method utilizes a stochastic elastic filling algorithm for coated particles to obtain the initial geometric parameter combination of fuel elements; constructs a multiphysics-coupled finite element model, and optimizes the final fuel design parameters under multiple constraints; based on this, Monte Carlo neutronics calculations are used to perform criticality optimization on the core, resulting in a neutronics design scheme that meets criticality and safety requirements; based on this core geometry, computational fluid dynamics analysis is used to optimize coolant distribution by adjusting the porosity of the thermally porous plate, and under conditions that meet temperature and pressure drop limits, the core power and hydrogen mass flow rate range that can achieve preset specific impulse, thrust, and thrust-to-weight ratio performance indicators are determined. Compared with existing technologies, this invention achieves a complete and systematic optimization design for a novel spherical metal-ceramic fuel nuclear thermal propulsion pebble bed reactor core, from concept to meeting preset propulsion performance indicators.

The invention belongs to the field of engines, and particularly relates to a fuel supply structure and a solid-gel hybrid rocket engine, the fuel supply structure comprises an oxygen-enriched grain propellant supply mechanism and a gel fuel supply mechanism; the oxygen-enriched grain propellant supply mechanism comprises a coating column, the coating column comprises a coating sleeve with an opening facing the combustion chamber, and an oxygen-enriched grain is arranged in the coating sleeve; the gel fuel supply mechanism comprises a gel fuel supply device, an annular channel and an atomizer which are sequentially communicated, the annular channel is wound outside the wrapping sleeve, the downstream of the annular channel protrudes out of the downstream of the wrapping sleeve, the atomizer is arranged at the end, protruding out of the downstream of the wrapping sleeve, of the annular channel, and the annular channel and the wrapping sleeve are combined to block the upstream end of the combustion chamber. The engine adopts a new combination of the gel fuel and the oxygen-enriched grain, has the advantages of the gel fuel and the oxygen-enriched grain, and achieves the effects of increasing specific impulse, adjusting thrust and starting for multiple times.

The invention relates to an air-breathing electric propulsion system ground test system and method, and belongs to the field of ultra-low orbit space electric propulsion, and the system comprises a spiral wave plasma source, an incoming flow parameter measurement device, a tested air-breathing electric propulsion system, a micro-thrust frame, a vacuum system, a gas cylinder and a micro-thrust frame measurement system. A gas cylinder is used as a gas source to generate simulated incoming flow through a spiral wave plasma source, the simulated incoming flow enters a tested air-breathing electric propulsion system, the thrust of the tested air-breathing electric propulsion system is calculated by measuring net thrust, and then specific impulse is calculated. The spiral wave plasma source is used for generating rarefied gas plasma incoming flow equivalent to neutral gas incoming flow, incoming flow with different track heights and different component contents can be simulated by adjusting flow, power, magnetic field intensity and the like, and performance verification of the air suction type electric propulsion system is achieved by measuring net thrust.

The application provides a metal fuel propellant and a preparation method thereof, and belongs to the technical field of solid propellants. 3 Compared with traditional aluminum-based propellants, the zirconium-based propellant has a greater density, and the increase of the density can not only increase the volume specific impulse of the propellant, but also increase the speed increment of a missile, so that the zirconium-based propellant has more excellent energy characteristics and can be applied to small volume engines; the application of zirconium hydride as a fuel additive increases the hydrogen content in the fuel, reduces the flame temperature of the propellant and improves the specific impulse of the propellant; the application of zirconium hydride as a fuel additive has good resistance to direct oxidation of perchlorate and can improve the energy characteristics and combustion performance of the zirconium-based propellant; and the application of aluminum powder as a fuel additive has high combustion heat value and good oxidation performance and can promote the combustion reaction of the zirconium powder.

The application relates to the field of aerospace propulsion technology, in particular to a space micro-thrust engine, which comprises an external supporting cylinder, a grid assembly is fixedly connected to one side of the external supporting cylinder, a flange plate is fixedly connected to the other side of the external supporting cylinder, a hollow cathode assembly is fixedly connected to the middle of the flange plate, the hollow cathode assembly is located in the external supporting cylinder, an annular gas pipe is arranged between the grid assembly and the hollow cathode assembly, one end of the annular gas pipe penetrates through the flange plate, a magnetic field part is arranged on the external supporting cylinder, and the magnetic field part is located at the end far from the flange plate. The application can provide a compact structure and high specific impulse space micro-thrust engine.

This invention discloses a combined power cycle system and method for cooling incoming air using ammonia-water, relating to the aerospace field. The system can switch between turbine-ramjet operating modes through an adjustable structure. Particularly during turbine-ramjet switching, ammonia-water is used to cool the high-temperature incoming air in the intake duct. The absorbed ammonia-water is then distilled to separate high-temperature ammonia gas and water. This invention combines heat exchanger precooling and jet precooling, effectively increasing the pressure ratio of the fan and compressor through multiple air cooling processes. This improves the compressor outlet gas temperature, thereby reducing compressor power consumption and providing greater thrust. Ammonia, as a supplementary fuel, has a higher specific impulse, addressing the thrust gap during engine mode transitions and effectively widening the operating speed range of the turbine engine.

The invention discloses a dual-fuel cross-medium power system and a control method thereof. The dual-fuel cross-medium power system comprises a combustion chamber, at least one group of inflow pipelines for conveying an air medium or a water medium into the combustion chamber, and a dual-fuel storage tank. The dual-fuel storage tank is fixedly assembled at the front end of the combustion chamber and provided with a hollow annular columnar shell, a grain assembly, an annular piston, an annular multi-point powder injection assembly and a fuel spray pipe. The front end of the shell is closed, the rear end of the shell is open, the shell comprises an inner shell and an outer shell, the grain assembly is fixedly assembled on the inner side of the inner shell in the axis direction of the inner shell, the inner shell is sleeved with the annular piston, the annular piston can slide in the axis direction of the inner shell, the outer side wall of the annular piston abuts against the inner wall of the outer shell, and a cavity between the outer shell and the inner shell is a powder fuel storage cavity. The front end of the annular multi-point powder injection assembly communicates with the powder fuel storage cavity, and the rear end of the annular multi-point powder injection assembly communicates with the combustion chamber. The power system is small in redundant mass and high in specific impulse.

The invention belongs to the technical field of power systems, and particularly relates to a quantum modified energy storage microsphere working medium turbofan power system and a laser explosion method. Aiming at the pain points of traditional turbofan power fuel oil pollution, low ignition precision and low energy efficiency, quantum modified bio-based solid microspheres with the intrinsic energy density being 7-8 MJ / kg and the diameter being 50 microns serve as a core working medium, and the quantum modified bio-based solid microspheres with the intrinsic energy density being 7-8 MJ / kg and the diameter being 50 microns are subjected to precise high-frequency impact through a 16-path independent VCSEL femtosecond pulse laser array (a VCSEL unit is 100 mW, a subunit is 1.35 W, 16 paths are 472 W, the frequency is 100 Hz, the electro-optical conversion efficiency is larger than or equal to 70%, the pulse width is 100-300 fs, the wavelength is 810 nm, and the focusing power density is 1 * 10 < 9 >-5 * 10 < 9 > W / cm < 2 >); the energy-releasing broken particles, atomized water drops and filtered air form a mixed working medium, and the intelligent control unit and the turbofan mechanism are combined to do work; the power is amplified by 22 times, 1% of the laser power causes a 13N / kW impulse coupling coefficient and 400-1200 s specific impulse, and the system is superior to a fuel oil power system. The technology is novel, low-power input is accumulated into high-power output, carbon emission is low, and the system is suitable for low-altitude economic traffic and high in popularization value.

The invention relates to a long-life and high-specific impulse hydrazine arc thruster, which relates to the field of electric thrusters, and comprises a rear shell, a front shell, a gas interface, an electric connector, a belleville spring, a cyclone, a cathode and an anode, the rear shell and the front shell are made of different materials and are fixedly connected through welding; the fuel gas interface is fixedly connected to the rear shell so as to introduce gas which is formed by catalytic decomposition of liquid anhydrous hydrazine into the rear shell; the electric connector is fixedly connected to the end part of the rear shell so as to supply power to the thruster; a plurality of belleville springs are arranged in the rear shell close to the electric connector; the swirler is installed in the front shell so as to convert gas input by the gas connector into circumferential rotational flow, the anode is welded to the front shell, the anode is made of tungsten-rhenium alloy which is low in rhenium content and doped with hafnium carbide, a throat is formed in the anode and integrated with the spray pipe, the sharp end of the cathode extends to the throat, arc discharge is generated between the sharp end of the cathode and the anode, and gas is heated and ionized. The invention has the advantages that the specific impulse can reach 600s, and the service life can reach 1000h.

Time of flight mass spectrometer device for characterizing different angle performance of electrospray thruster belongs to the field of test of launch performance of electrospray thruster.The present application comprises a time of flight mass spectrometer device, a thruster rotating system, a thruster, a power supply system and a data acquisition device.The time of flight mass spectrometer device comprises an electrostatic gate device, a particle flight tube section and a collection device.The thruster device comprises an electrospray thruster, an electrospray thruster rotating system and a fixing device.The collection device is a stainless steel circular tube with a flange, and an insulating gasket, a metal grid and a collection plate are arranged in the circular tube.The present application wraps the transmission line for transmitting the current signal on the transmission collection plate with a metal mesh to reduce electromagnetic interference.The launch angle of the electrospray thruster is controlled by an electric rotating table outside the vacuum chamber, the rotating angle of the thruster is accurately controlled, the plume velocity of the electrospray thruster under different angles is measured, and the thrust and specific impulse of the electrospray thruster under different angles are calculated through the measured data.

The application relates to the technical field of aerospace propulsion, and particularly discloses a vacuum arc pulse plasma thruster. The vacuum arc pulse plasma thruster comprises a vacuum arc ignition system and a pulse plasma acceleration system. The vacuum arc ignition system forms an arc to generate plasma based on high-voltage pulse discharge ablation of metal, and injects the plasma into the pulse plasma acceleration system; the pulse plasma acceleration system axially accelerates the plasma based on the Lorentz force generated by the internal formed orthogonal electromagnetic field, and the high-speed ejection of the plasma from the discharge channel forms a recoil force. The introduction of the pulse plasma propulsion acceleration mechanism promotes the efficient conversion of the capacitor energy storage into axial directional kinetic energy, and improves the thrust, specific impulse and energy conversion efficiency of the thruster.

The application discloses a gas generator of an air turbine rocket engine, which comprises a shell, an ignition device, a liquid supply system and a combustion chamber; the gas generating agent used by the gas generator comprises ADN ionic liquid solution, kerosene and high-metal solid fuel; the liquid supply system comprises an ionic liquid storage tank and a kerosene storage tank, and the ionic liquid storage tank and the kerosene storage tank store ADN ionic liquid solution and kerosene respectively; the liquid supply system adopts an electric pump to supply liquid to the combustion chamber in the shell, and the electric pump is used to adjust the delivery pressure and flow rate; the ignition device is arranged in the combustion chamber, and a high-metal solid fuel grain is arranged at the tail of the inner cavity of the shell; the high-metal solid fuel grain is of an inner hole combustion type, and the surface of the central hole is of a spiral groove structure. The application has the advantages of solid high energy density and liquid adjustability, overcomes the problems of high danger of traditional hydrogen peroxide and low-temperature use of liquid oxygen, and significantly improves the specific impulse, safety and mission adaptability of the engine.

The embodiment of the invention discloses an electric thruster specific impulse test system and method. The system comprises a double-gate probe, a metal shielding shell, a support and a detection module. The two ends of the metal shielding shell are respectively provided with a double-gate probe, and the metal shielding shell and the double-gate probes at the two ends form a closed cavity. The support is used for fixing the metal shielding shell and an electric thruster to be tested. The double-grid probe comprises two grid electrodes which are arranged in parallel, and micropores which are distributed in an array are formed in the grid electrodes; the detection module is electrically connected with the double-gate probes at the two ends of the metal shielding shell and used for measuring the time difference of plume particles of the electric thruster passing through the double-gate probes at the two ends of the metal shielding shell; the detection module internally comprises a capacitor connected with the double-gate probe in parallel, and when plume particles of the electric thruster pass through the double-gate probe, the capacitor discharges. According to the technical scheme provided by the invention, the problems of equipment complexity and relatively low precision and reliability are relieved; in addition, by measuring discharge of the capacitor, stable and high-intensity current can be obtained, and reliable measurement is easy to carry out.

The application discloses a kind of remaining carrying capacity evaluation methods based on numerical programming after specific impulse drop, use failure occurrence time, remaining thrust percentage and propellant leakage percentage as the characteristic quantity of characterizing specific impulse drop failure.By searching the fault adaptation range of rocket into target orbit and safe parking orbit under different characteristic quantity combination conditions offline, the position that the remaining carrying capacity after failure can send the load to is evaluated.Finally, the fault adaptation range is bound to the rocket offline, and according to the three quantities of failure occurrence time, remaining thrust percentage and propellant leakage percentage, the corresponding adaptation range of the failure is quickly found after the failure occurs, and the remaining carrying capacity evaluation is realized.

The application discloses a kind of liquid rocket engine thrust keeping multivariable fault-tolerant control methods, comprising: analyzing liquid rocket engine failure mechanism and influence characteristics, establish typical fault model;With mixing ratio as constraint quantity, construct multivariable predictive optimal controller for thrust keeping, and design fault-tolerant control architecture verifies the performance of control logic under typical fault injection.The application takes the combustion chamber pressure reflecting thrust as the controlled variable, the mixing ratio as the constraint quantity, the flow regulator valve core and the fuel main valve opening as the control variable, based on the neural network prediction model offline training, and according to the error containing penalty function, the control variable quadratic form index is optimized on-line, the control variable for thrust keeping is calculated, and the maximum change of mixing ratio is small, and the specific impulse is higher.The application reduces the transient loss and steady loss of engine after failure, and meets the demand of thrust keeping and mixing ratio constraint in main stage working condition, and has strong fault-tolerant capability.

This invention relates to the field of space electric propulsion technology, and particularly to a method for refined thruster allocation in an electric propulsion system based on thrust vector parameters. The method includes: for each electric thruster, determining the first and second angles between the thrust vector of the current electric thruster and the coordinate system of the entire satellite when the thrust vector points towards the center of mass of the entire satellite, based on the thrust vector parameters of the current electric thruster and the transformation model between the thrust vector of the electric thruster and the coordinate system of the entire satellite; determining the propellant consumption for maintaining north-south position using the current electric thruster, based on the specific impulse of the current electric thruster and the first and second angles between the thrust vector of the current electric thruster and the coordinate system of the entire satellite when the thrust vector points towards the center of mass of the entire satellite; and sorting all the electric thrusters to be allocated according to their propellant consumption based on the propellant consumption of each electric thruster to identify and allocate the target electric thrusters, thereby reducing the propellant load of the electric propulsion system and improving the overall payload capacity of the satellite.

The invention provides a dual-mode Hall and cathode thruster space electric propulsion system and a working method.The dual-mode Hall and cathode thruster space electric propulsion system comprises a Hall thruster, a cathode thruster, a storage box, a power system and a pipeline flow control device system, and the storage box is connected with the Hall thruster and the cathode thruster through the pipeline flow control device system; the pipeline flow control system is used for gas supply switching and flow control of a pipeline, the power supply system is connected with the Hall thruster and the cathode thruster through the power supply switching switch, and the power supply system is used for power supply switching. By sharing the storage and supply system and the power supply system, the size and weight of the propulsion system can be greatly reduced, high specific impulse and high thrust dual-mode work is realized, and the requirements of a multifunctional, lightweight and small-size power system of a microsatellite can be met; by adopting the dual-mode work, the risk hidden danger caused by independently applying the cathode thruster to execute the flight task is obviously reduced.

The application provides a dual-mode solid working substance micro electrothermal thruster device, which comprises an anode nozzle, a capillary cavity, a trigger electrode and a cathode base; the trigger electrode and the cathode base have a gap therebetween, and the gap constitutes a trigger surface, which is used for generating initial plasma through surface discharge and inducing main gap discharge in the capillary cavity to generate plasma; after the main gap discharge is formed, different voltage forms are applied to make the solid working substance discharge loop of the capillary cavity conductive, so that the plasma continuously moves to the anode nozzle to generate thrust through external ejection. The dual-mode solid working substance micro electrothermal thruster provided by the application can realize two working mode outputs of “large thrust and low specific impulse” and “high specific impulse and small thrust”, has the advantages of simple structure, low mass, low system power consumption, high working substance utilization rate and wide output parameter range, and is suitable for more application scenarios.

The present application relates to a kind of on-orbit engine thrust calibration method, device, electronic equipment and storage medium.The method comprises: obtaining the theoretical output of current time on-orbit engine and the real orbit change result of current time orbit change;Theoretical output includes theoretical thrust and theoretical specific impulse, and real orbit change result includes real semi-major axis and real orbit inclination;According to theoretical output, the theoretical orbit change result of current time orbit change is calculated, and theoretical orbit change result includes theoretical semi-major axis and theoretical orbit inclination;Based on the deviation of real orbit change result and theoretical orbit change result, theoretical output is cyclically corrected until the final correction value of theoretical output is obtained;Deviation includes the semi-major axis deviation between theoretical semi-major axis and real semi-major axis, and the inclination deviation between theoretical orbit inclination and real orbit inclination;The final correction value of theoretical output is used as the thrust calibration result of current time on-orbit engine orbit change.This scheme can independently calibrate the thrust of on-orbit engine, and the accuracy of calibration result is higher.

The invention belongs to the field of aerospace, and particularly relates to a fuel weight coefficient estimation method for a climbing-type aerospace vehicle, and the method comprises the steps: obtaining a fuel type of the climbing-type aerospace vehicle; the gravitational acceleration at the corresponding position of the ground serves as the gravitational acceleration during climbing, and the corresponding relation of the fuel coefficient of the climbing aerospace vehicle, the rocket power specific impulse and the flight-launch integrated design efficiency is calculated; according to the rocket power specific impulse and the flight-launch integrated design efficiency, the rocket power effective specific impulse is determined through multiplication; the method can accurately calculate the effective specific impulse of rocket power, and is suitable for a concept scheme design stage of a cross-atmosphere climbing aerospace vehicle. Calculation can be completed through simple mathematical formulas and design parameters, and tedious calculation processes and data support requirements are reduced. Meanwhile, the method can improve the accuracy and efficiency of estimation, and provides powerful support for optimization of aircraft design and performance evaluation.