120 results about "Air breathing" patented technology

The invention relates to a propulsion system, in particular to an air-breathing radio frequency ion electric propulsion system with mixed working medium power and a multi-mode working method thereof, and aims to solve the problems that an existing air-breathing electric propulsion system only takes atmosphere as a single working medium, the atmosphere trapping efficiency is low, and the performance of a spacecraft is affected. The device comprises a propelling and gas trapping unit, a radio frequency neutralizer, a power and control unit and a xenon storage and supply unit, the propelling and gas trapping unit comprises a radio frequency ion thruster and an atmosphere trapping gas inlet channel; the power and control unit comprises a direct current module, a control module and a radio frequency module; the xenon storage and supply unit comprises a xenon bottle and a valve and sensor assembly; the control module and the valve are electrically connected with the sensor assembly. The radio frequency ion thruster is electrically connected with the radio frequency neutralizer through the radio frequency module. The first output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency neutralizer, and the second output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency ion thruster.

The invention belongs to the technical field of aero-engines, and particularly relates to an aero-engine air inlet flow pipe calibration method, which comprises the following steps of: measuring internal flow field parameters in a standard flow measurement section through a plurality of sensors, and calculating theoretical air mass flow qm1TH of the standard flow measurement section; the actual air mass flow qm1ACT of the standard flow measuring section is measured through a negative pressure air suction type flow standard device; the outflow coefficient of the standard flow measurement section is calculated; connecting a calibrated engine air inlet flow pipe A2 to the front of the standard flow measuring section, and connecting the engine to the rear of the standard flow measuring section; calculating the actual air mass flow of the standard flow measurement section, and taking the actual air mass flow as the actual air mass flow qm4ACT of the calibrated engine air inlet flow pipe A2; and the outflow coefficient of the calibrated engine air inlet flow pipe A2 is calculated based on the theoretical air mass flow qm4TH and the actual air mass flow qm4ACT of the calibrated engine air inlet flow pipe A2.

The invention provides an adjustable air-breathing power ultra-low orbit satellite configuration. The adjustable air-breathing power ultra-low orbit satellite configuration comprises an adjustable closed head part, a body, a stable wing surface and a contraction tail part, the adjustable closing head and the contraction tail are connected to the two opposite ends of the body respectively, and the stable wing face is connected to the side, close to the contraction tail, of the body. The design that the adjustable closing head is an openable / closable head is adopted, the adjustable closing head is a spherical head after being closed when a satellite does not need to inhale, and resistance is reduced; and after being expanded, the air inlet channel can be contracted into an inwards-concave air inlet channel, so that active compression of air is achieved. Therefore, task requirements under different flight heights can be met through flexible switching of the inspiration working mode and the non-inspiration working mode, the aerodynamic resistance of the satellite in an ultra-low orbit is effectively reduced, and the hang time of the satellite is prolonged.

The invention discloses a rocket-based rotary detonation engine combustion chamber structure and application thereof, and belongs to the technical field of rocket engine research. The combustion chamber structure comprises an isolation section, an equal-direction oil injection section and an expansion-concave cavity section which are connected in sequence. A central rocket is arranged in the isolation section and the equal direct oil injection section; and four groups of first ring-jet injection points are arranged on the equal straight oil injection section, two groups of second ring-jet injection points are arranged on the central rocket, and two groups of third ring-jet injection points are arranged on the expansion-concave cavity section. Oil is injected into the expansion-concave cavity section firstly, a pre-detonation pipe is adopted for igniting an oil-gas mixture, rotating detonation waves are formed, the built-in center rocket can provide thrust for the aircraft at the ground zero speed and form an injection effect for air suction, extra thrust is provided when the cruising thrust of the aircraft is insufficient, and when the rotating detonation combustion flameout boundary exists, the detonation effect is achieved. A flame stabilizing mechanism is provided for rotary detonation combustion, and stable work of wide-speed-range and wide-height-range rotary detonation combustion is achieved.

The invention provides an air suction type electric propulsion air suction chamber. The air suction type electric propulsion air suction chamber comprises grids, an air inlet channel and a rectifying device. The bottom of the grid is connected with the front end of the air inlet channel, the inner wall of the air inlet channel is of a parabola-shaped structure, the rear end of the air inlet channel is connected with the long pipe, the rectifying device is installed in the air inlet channel through the supporting frame, the rectifying device is of a convex-mirror-like structure, the inner wall of the rectifying device is of a parabola-shaped structure, and the focus of the rectifying device is the same as the focus of the parabola-shaped structure of the inner wall of the air inlet channel in position. The opening direction of the rectifying device is opposite to the opening direction of the air inlet channel. According to the invention, the incoming flow gas is refracted twice and is finally injected into the long pipe at the rear end of the air suction chamber in parallel, so that collision and corrosion of air particles to the wall surface of the long pipe are reduced, the service life and the performance of the air suction type electric propulsion are greatly improved, the gas collection rate is improved, and a new technical approach is provided for the development of the air suction type electric propulsion.

The invention provides a trajectory online planning method for an air-breathing hypersonic aircraft with an unstarted air inlet, and the method comprises the steps: constructing a mechanism model of the air-breathing hypersonic aircraft, and correcting the mechanism model according to the output of a neural network and in combination with an air inlet / isolation section pressure sensor array; and in the flight process, stability margin prediction is carried out based on the corrected mechanism model, and when the air inlet non-starting margin is lower than a warning value in the future, the trajectory optimization model is called to carry out online trajectory planning. The output error of the corrected mechanism model is obviously reduced, and more accurate dynamic characteristic data supports reduction of the error between the control effect and the preset index; the accuracy of the mechanism model is obviously improved; once it is found that the non-starting margin of the air inlet channel is lower than the warning value in the future, the track optimization model is called to conduct online track planning, and therefore the situation that the air inlet channel is not started is avoided; the method has remarkable effectiveness and robustness in the aspects of trajectory optimization and air inlet non-starting protection.

An air breathing engine system has an engine inlet configured to receive a flow of air. A combustor is in fluid communication with the engine inlet. A rocket ejector is positioned between the engine inlet and combustor. The rocket ejector is connected to the combustor with a mixing tube. The rocket ejector adjustably outputs high-temperature gas into air in the mixing tube. The high-temperature gas changes a pressure and thermal energy of the air within the mixing tube, which can be used to optimize combustor operation, such as based on a trajectory, flight condition, or speed of an aircraft propelled by the air breathing engine system.

The invention discloses a guide wall for airplane ground test run, which comprises a U-shaped barrier part enclosing an airplane, the barrier part comprises a side guide wall corresponding to the side of the airplane and a rear guide wall corresponding to the tail of the airplane, the side guide wall comprises a windward side and a leeward side, and the leeward side is connected with a fixing frame. The fixing frame is connected with a ground anchor bolt to support the side flow guide wall, and the rear flow guide wall is fixedly connected with ground concrete; the length of the side flow guide wall exceeds the position of the nose during the test run of the airplane, and an air inlet channel is arranged at the side flow guide wall corresponding to the side wing during the test run of the airplane, penetrates through the side flow guide wall and supplements airflow to the engine close to the side wing of the air inlet channel; by arranging the air inlet channel, air flow is introduced from the side face of the aircraft directly facing the engines, the problem that the test is affected by insufficient air suction of the double engines due to the fact that the side face flow guide wall is far longer than the fuselage is solved, the internal and external pressure difference of the flow guide wall is reduced, and test run of the aircraft is facilitated.

The invention relates to the technical field of aircrafts, in particular to an air-breathing power system and method for a Mars surface aircraft, and the air-breathing power system comprises an air inlet channel, an isolation section, a combustion chamber, a cooling channel, a spray pipe, a heat exchanger and a fuel device, the fuel device comprises a liquid fuel tank, a high-pressure pump, a magnesium powder storage tank, a powder feeder and a gasification nozzle. The air-breathing power system of the aircraft is designed aiming at the fact that gas on the surface of Mars is mainly composed of CO2, the air-breathing power system is provided with magnesium powder capable of reacting with CO2 and methane or hydrogen carrying the magnesium powder to be sprayed into a combustion chamber, CO2 needed by combustion is provided through compressed air in the high-speed flight process, and the combustion efficiency is improved. The air-breathing type power system does not need to carry a large amount of fuel and oxidizing agents like a common rocket engine, and the specific impulse of the engine of the air-breathing type power system is far higher than that of the engine of a common rocket.

A vessel for transporting cargo by water comprises one or more propulsion devices with at least one air-breathing engine from the one or more propulsion devices disposed entirely above a surface of the water in a high-speed mode of operation. The least one air-breathing engine is configured to bleed airflow into a fluid circuit. The fluid circuit routes the bleed airflow to one or more fluid outlets disposed entirely below the surface of the water in the high-speed mode of operation. At least some of the one or more fluid outlets release the bleed airflow in the high-speed mode of operation generate a boundary layer along at least a portion of a hull of the vessel to reduce a skin friction drag thereof.

The present invention relates to a spacecraft electric thruster particularly suitable for both propellant-fed and air-breathing electric space propulsion comprising an ionization chamber (11) for plasma generation which develops along a central development axis (X), wherein the ionization chamber (11) has a chamber inlet (12) configured to allow gas inflow and a chamber outlet (13) configured to allow extraction of ionized gas particles; a power injection assembly (30) configured to perform electron heating for plasma generation inside the ionization chamber (11); a magnetic assembly (20) configured to generate a magnetic field for plasma confinement in the ionization chamber (11); and an ion extraction and acceleration assembly (40) configured to extract ionized gas particles from the ionization chamber (11) through the chamber outlet (13) and accelerate them along an ion acceleration direction (Z), wherein the power injection assembly (30) comprises a microwave power delivery system (31) configured to deliver an electric field oscillating at a selected microwave frequency (f) and to inject the generated oscillating electric field in the ionization chamber (11), and wherein the ionization chamber (11) comprises a plurality of walls (11a, 15,16,41) closed to the passage of microwaves apart from a portion of at least one lateral wall (11a) of the plurality of walls (11a, 15,16,41) at which the power delivery system (31) is coupled to inject the generated oscillating electric field in the ionization chamber (11), and is resonant for electromagnetic fields oscillating at the selected microwave frequency.

The application provides a kind of air-breathing hypersonic vehicle aerodynamic servo elasticity active control system and method, comprising: elasticity frequency estimator obtains the elasticity frequency estimation value of controlled object based on the pitch angle information sent by controlled object, and sends to state disturbance joint estimator;State disturbance joint estimator obtains corresponding state estimation result, disturbance estimation result and pitch angle estimation result based on the pitch angle information received, elevator deflection angle and elasticity frequency estimation value;Attitude controller obtains the elevator deflection angle of controlled object based on the rigid body pitch angle control value received, state estimation result and disturbance estimation result, and sends to state disturbance joint estimator and controlled object.The application can control the stability of controlled object.

The invention relates to an air suction type cusped field electric propulsion system with multiple working medium supply modes and a supply method. The invention relates to the technical field of air suction type electric propulsion systems. The system comprises an electric thruster and collection device module, a microwave cathode module, a xenon storage and supply module and a control and energy supply module. According to the air-breathing cusped field electric propulsion system with the multiple working medium supply modes, the adaptive capacity of the electric thruster under the complex working condition is improved through the xenon working medium carried by the air-breathing cusped field electric propulsion system while the ultra-low orbit rarefied atmospheric environment is fully utilized. In addition to an ultra-low rail air-breathing type electric propulsion task, the xenon working medium carried by the ultra-low rail air-breathing type electric propulsion device can enable the ultra-low rail air-breathing type electric propulsion device to give consideration to a higher rail range, and a traditional xenon electric propulsion rail interval and an ultra-low rail space of 150 km or above are covered.

The invention relates to the technical field of low-orbit aircrafts, in particular to an airfoil auxiliary control air-breathing electric propulsion upper-layer atmosphere aircraft which comprises a main aircraft body, a deflection folding assembly, an electric driving system and a power supply system. The deflection folding assembly is used for locking the position of the folding wing folded on the main wing before the aircraft takes off, and is used for actively unlocking the main wing to unfold the folding wing and locking the position of the unfolded folding wing after the aircraft enters an orbit; when the folding wings of the aircraft are in the unfolded state, the folding wings are driven to deflect so as to adjust the deflection angles of the folding wings relative to the surfaces of the main wings; the power supply system is used for supplying power to the electric driving system and the deflection folding assembly. According to the device, the folding wings are folded to reduce the size of the whole aircraft, so that the requirement on the space size of the carrier is reduced; the aerodynamic lift-drag ratio of the aircraft is adjusted by adjusting the deflection angle of the folding wings, and then aerodynamic force control capacity is provided for the aircraft.

The invention provides an air-breathing engine incremental control method based on reinforcement learning, and relates to the technical field of air-breathing engine control. The deep reinforcement learning control module is used for outputting control increments of control variables such as the main fuel flow, the inner culvert and outer culvert thrust augmentation fuel flow and the nozzle throat area, the safety limiting module is combined for monitoring and cutting multiple monitoring parameters in real time, and finally a smooth and feasible execution signal is generated through the action space processing module. And multi-dimensional fine adjustment of the engine is realized. According to the scheme, multi-agent parallel training is adopted, switching between a low thrust mode and a high thrust mode is supported, under the condition that an accurate mathematical model is not needed, nonlinearity and coupling characteristics of an engine are flexibly coped, one-time large jump is prevented, safety and fuel efficiency are improved, and the requirements for high thrust and low fuel consumption in the flight process are met.

The invention discloses a digital twinning system of an air-breathing strong precooling engine test bed, the system comprises a real-time simulation part and a performance prediction part, the real-time simulation part carries out real-time analogue simulation on the test process through modular modeling, steady-state / dynamic calculation and physical test bed data coupling; according to the performance prediction sub-part, a reference model and an actual state model are established based on a long-short-term memory neural network, and performance prediction is conducted on the performance of the strong pre-cooling engine. The digital twin system provided by the invention is convenient for variable working condition adjustment, parameter adjustment boundary exploration and error risk and debugging cost reduction, so that a debugging rule can be summarized in a test bed debugging process, a debugging scheme of a physical test bed is assisted to be formulated, and a safe and efficient test process is ensured.

The invention provides an air-breathing aircraft for long-term flight in an upper atmosphere, and relates to the technical field of aircrafts in the upper atmosphere, the air-breathing aircraft comprises an aircraft body, the aircraft body is provided with an air-breathing electric propulsion system, the outer contour of the aircraft body is cylindrical, and an air inlet channel is arranged in the center of the aircraft body; the air inlet channel comprises a gradually-shrinking conical inlet pipeline, an inwards-concave gradually-shrinking compression pipeline, a cylindrical ionization acceleration pipeline and a gradually-expanding conical outlet pipeline which are sequentially arranged from front to back. The solar cell wings are arranged on the outer side of the aircraft body, the solar cell wings are forward-swept trapezoidal wings, and the solar cell wings can supply power to the air suction type electric propulsion system; the problem that in the prior art, technical defects exist in the aerodynamic configuration design of how to achieve thrust and resistance balance of an aircraft flying for a long time in the upper atmosphere is solved.

A self-contained vertical takeoff and landing vehicle (VTOL) with aspect ratio of height to width no greater than 5.3:1 and with both rocket and airbreathing propulsion systems that can selectively operate one at a time or both at the same time and fuel tanks that have internal lattice structures formed of groups of interconnected multiplicities of fuel cells such as gyroids that line and are integrated with a vehicle skin to form a structural skeleton strengthened with said lattice structures that also serve to reduce fuel slosh compared with fuel tanks lacking such lattice structures and provide other benefits such as fuel temperature profile control.

The invention provides a pneumatic servo elastic active control system and method for an air-breathing hypersonic aircraft, and the method comprises the steps: an elastic frequency estimator obtains an elastic frequency estimation value of a controlled object based on the pitch angle information transmitted by the controlled object, and transmits the elastic frequency estimation value to a state disturbance joint estimator; the state disturbance joint estimator obtains a corresponding state estimation result, a disturbance estimation result and a pitch angle estimation result based on the received pitch angle information, elevator deflection angle and elastic frequency estimation value; and the attitude controller obtains the elevator deflection angle of the controlled object based on the received rigid body pitch angle control value, the state estimation result and the disturbance estimation result, and sends the elevator deflection angle to the state-disturbance joint estimator and the controlled object. According to the invention, the controlled object can be stably controlled.

Ejector for use in a refrigeration circuit device, which features: a drive-side nozzle (31) designed to expand a drive-side coolant and accelerate it to a supersonic speed and inject the drive-side coolant as a drive-side injection coolant in a gas-liquid two-phase state; and a body (32) which includes: a coolant suction port (321) designed to draw suction-side coolant; a suction-side nozzle (322) designed to expand and inject the suction-side coolant, which is drawn in as suction-side injection coolant through the coolant suction port (321); a mixing section (323) designed to mix the suction-side injection coolant injected from the suction-side nozzle (322) and the drive-side injection coolant injected from the drive-side nozzle (31); and a diffuser section (324) designed to convert kinetic energy of the mixed coolant formed from the suction-side injection coolant and the drive-side injection coolant mixed in the mixing section (323) into pressure energy, wherein: the drive-side nozzle (31) has a converging section (31b), a neck section (31c) and a diverging section (31d), wherein the diverging section (31d) has a passage cross-sectional area for the passage of the coolant, which is progressively enlarged from the neck section (31c) to a drive-side injection port (31e); the drive-side nozzle (31) is arranged coaxially in a cylindrical space formed on an inside of the body (32); a coolant passage of the suction-side nozzle (322) is formed between an outer circumferential surface of the diverging section (31d) and an inner circumferential surface of the body (32) and has a cross-section that is formed in an annular shape; the drive-side injection port (31e) of the drive-side nozzle (31) and a suction-side injection port (322a) of the suction-side nozzle (322) are open such that an injection direction of the drive-side injection coolant injected by the drive-side injection port (31e) and an injection direction of the suction-side injection coolant injected by the suction-side injection port (322a) coincide; a pressure of the drive-side injection coolant is higher than a pressure of the suction-side injection coolant; and when the mixed coolant flows out of the mixing section (323) and into the diffuser section (324), a pressure of the mixed coolant measured on a central axis of the mixing section (323) and a pressure of the mixed coolant measured on a wall surface of the mixing section (323) are equal, while a flow velocity of the mixed coolant is a subsonic velocity.

A vessel for transporting cargo by water comprises one or more propulsion devices with at least one air-breathing engine from the one or more propulsion devices disposed entirely above a surface of the water in a high-speed mode of operation. The least one air-breathing engine is configured to bleed airflow into a fluid circuit. The fluid circuit routes the bleed airflow to one or more fluid outlets disposed entirely below the surface of the water in the high-speed mode of operation. At least some of the one or more fluid outlets release the bleed airflow in the high-speed mode of operation generate a boundary layer along at least a portion of a hull of the vessel to reduce a skin friction drag thereof.

The present application belongs to the field of overall design of single-stage orbital aerospace vehicles, and particularly relates to a method for estimating the mass of an aerospace vehicle. It includes: step one, obtaining the general equation of the vehicle acceleration process; step two, constructing a mass ratio estimation model of the rocket propulsion mode according to the general equation of the vehicle acceleration process; step three, constructing a mass ratio estimation model of the air-breathing propulsion mode according to the general equation of the vehicle acceleration process and the mass ratio estimation model of the rocket propulsion mode; step four, constructing a mass estimation model of the single-stage orbital aerospace vehicle according to the mass ratio estimation model of the rocket propulsion mode and the mass ratio estimation model of the air-breathing propulsion mode. The present application completes the estimation of the total mass of the single-stage orbital aerospace vehicle, and can quickly obtain the law between the power mode transition point, the vehicle structure mass fraction, the orbital load and the total mass of the vehicle, thereby improving the convergence speed of the solution.

This invention provides an air-breathing electric propulsion air chamber, including a grid, an air inlet, and a rectifying device. The bottom of the grid is connected to the front end of the air inlet, the inner wall of the air inlet has a parabolic structure, and the rear end of the air inlet is connected to a long tube. The rectifying device is mounted inside the air inlet via a support frame. The rectifying device has a convex mirror-like structure, with a parabolic inner wall. Its focal point is the same as the focal point of the parabolic structure on the inner wall of the air inlet, and the opening of the rectifying device faces in the opposite direction to the opening of the air inlet. This invention refracts the incoming gas twice, finally injecting it parallel into the long tube at the rear end of the air chamber. This reduces the collision and corrosion of the long tube wall by air particles, thereby greatly improving the lifespan and performance of the air-breathing electric propulsion system and increasing the gas collection rate, providing a new technical approach for the development of air-breathing electric propulsion.

The application provides a kind of collaborative jet airfoil optimization design method based on combined parameterization, belongs to the field of aircraft aerodynamic design and flow control.This method firstly carries out CST parameterization to the original basic airfoil for constructing original collaborative jet airfoil;With blowing / air suction port position, size as design variable, the curve of subsidence section is fitted by original basic airfoil linear interpolation, realizes the adaptive association of subsidence section and original basic airfoil.Subsequently, a joint design space containing CST parameterization variable, collaborative jet airfoil design variable and jet control parameter is constructed, the initial proxy model is constructed and updated to convergence by using SBMO framework.The application carries out global collaborative optimization to the subsidence section channel geometry and the overall profile of original basic airfoil, solves the parameter coupling problem of the basic geometry shape of collaborative jet airfoil and subsidence section, and can significantly improve the aerodynamic performance of collaborative jet airfoil.

The application discloses a combined engine and a control method thereof. The combined engine comprises a control assembly, an air inlet channel, a turbine core engine, a nuclear reactor system, a propellant supply system and a nozzle assembly which are connected through pipelines. The control assembly is used for controlling the air inlet end of the nuclear reactor system to be selectively connected with the air inlet channel, the turbine core engine or the propellant supply system through the pipelines. The air outlet end of the nuclear reactor system is connected with the nozzle assembly. The air inlet end of the nuclear reactor system is connected with the air inlet channel, the turbine core engine and the propellant supply system through the pipelines respectively by the control assembly, so that the combined engine enters the nuclear ram mode, the nuclear turbine mode and the non-suction mode respectively. The combined engine has the characteristics of high fuel economy, working in the atmosphere-free environment, realizing different speed domain flights by using different engine modes and meeting the needs of space transportation of the super-speed aircraft.

The application discloses an ultra-low orbit adsorption type air-breathing pressurization system, which is used for an ultra-low orbit electric propulsion system and comprises an adsorption pressurizer on one side and a gas storage on the other side; the adsorption pressurizer is used for adsorbing rare thin gas molecules under low-temperature conditions and releasing the gas molecules under high-temperature conditions, and generates a pressurization effect; and the gas storage is used for storing the pressurized gas molecules and releasing the gas molecules on demand. The application discards the traditional mechanical compression mode and innovatively adopts an adsorption type pressurization principle. Through integrated design of a rotating staggered valve and an adsorption plate group, gas guiding, flow channel switching and adsorption pressurization functions are integrated into one, forming a highly compact overall structure. The application discards the traditional mechanical compression mode and innovatively adopts the adsorption type pressurization principle. The problems of a complex structure, a large size, extremely high power consumption and prominent energy contradiction of the prior art ultra-low orbit electric propulsion system are solved.

This invention belongs to the field of plasma flow control technology for aircraft, specifically relating to an air-breathing plasma jet generator suitable for active flow control of aircraft. It is installed within an internal cavity of the aircraft, connected to an air intake and a nozzle. The air-breathing plasma jet generator generates orthogonal electric and magnetic fields. The electric field includes components in the direction of airflow and perpendicular to airflow. The magnetic field is used to confine electrons to gyroscopic motion and to ionize them through collisions with gas molecules. Furthermore, the magnetic field causes the ion current perpendicular to the airflow direction to generate an Ampere force in the direction of airflow under the influence of the magnetic field. This invention effectively solves the problem of exciter failure during high-speed flight by embedding the exciter entirely inside the aircraft and combining an innovative design of integrated excitation using magnetic fields, thermal energy, and acoustic waves, significantly improving the speed applicability range of plasma flow control technology.

The invention relates to the technical field of air intake systems of aero-engines, and discloses a self-adaptive hybrid air suction system of an aero-piston engine, which comprises an engine body for providing power for aviation flight, and is characterized in that an air suction unit for air intake is fixedly arranged at one end of an inner cavity of the engine body; the air suction unit comprises an air suction part used for air suction, a vortex generator used for improving air mixing and an air inlet adjusting mechanism used for adjusting the air inlet angle, the air inlet adjusting mechanism is fixedly connected with the air suction part through the vortex generator, and an air compression unit used for compressing air is fixedly arranged at one end of the air suction unit; the end, away from the air suction unit, of the air compression unit is fixedly connected with a combustion unit used for combusting compressed air, the end, away from the air compression unit, of the combustion unit is fixedly connected with an exhaust unit used for exhausting air to provide power, the air inlet mechanism can be adjusted to dynamically adapt to the flight working condition, the air inlet efficiency and stability are improved, and the response speed is increased; and the combustion efficiency is optimized.

The application provides a solar energy suction type electric propulsion unmanned aerial vehicle, which comprises a fuselage, wherein a ring truss type photoelectric conversion device is arranged on the fuselage, the photoelectric conversion device comprises a ring truss unfolding mechanism and a foldable thin film type photoelectric conversion array, the thin film type photoelectric conversion array is supported and connected by the ring truss unfolding mechanism, and the unfolding or folding of the thin film type photoelectric conversion array is realized through the ring truss unfolding mechanism; a thrust device is arranged on the fuselage, and the thrust device adopts a suction type electric propulsion device. The application uses the ring truss type photoelectric conversion device, can realize a large deployment and folding ratio design of the solar energy thin film, has a simple structure and can effectively reduce the structural mass. The application uses the suction type electric propulsion as a power device, does not need to carry additional propellants, can take gas from the surrounding thin atmosphere as the propellants, and can generate corresponding thrust after the gas is ionized and accelerated, so that the unmanned aerial vehicle can work on the orbit for a long time.

The invention discloses a recognition method and device for the working state of an air suction type liquid engine and a control system. The recognition method comprises the steps that power characterization signals of an engine are collected, and the power characterization signals comprise at least one of an optical signal of a combustion chamber of the engine, a pressure signal of the combustion chamber, an acceleration signal of an aircraft and a separation signal of the aircraft; and judging whether the engine is shut down abnormally or not based on the collection result of the power representation signal so as to recognize the working state of the engine. Therefore, the working state of the engine is recognized.