104 results about "Thrust chamber" patented technology

The invention discloses a multi-scale bidirectional coupling sweating cooling simulation method which comprises the following steps: firstly, on a mesoscale, constructing a mesoscopic flow model, simulating flow of a coolant, and obtaining parameters such as porosity, permeability and inertia coefficient of a porous medium; introducing the obtained parameters into a porous medium panel area of a thrust chamber geometric model constructed under a macroscale to realize cross-scale parameter coupling; the flow and heat transfer process in the thrust chamber geometric model after cross-scale parameter coupling is simulated, and the speed, pressure, temperature and local heat flux density distribution of a combustion chamber and a porous medium panel area under the macroscale are obtained; and taking the local heat flow density as a new wall surface boundary condition, capturing the flow and heat transfer characteristics of a coolant in pores under a thermal state working condition based on the mesoscopic flow model, and further evaluating the overall cooling performance and the thermal protection effect of the porous medium panel. Through coupling calculation of the mesoscopic pores and the macroscopic combustion chamber, the accuracy and reliability of prediction of the cooling performance of the porous medium are improved.

The invention provides a multi-channel temperature parameter pre-cooling process liquid rocket engine temperature field display method, and solves the problems that in the prior art, as a data acquisition system measures temperature parameters of a plurality of measuring points when a power system carries out preparatory work, the temperature parameters of the key temperature measuring points are measured on the basis of a temperature sensor; and the change condition of a structural system cannot be visually displayed after long-time numerical value display. According to the method, on the basis that simulation calculation is high in time complexity and cannot achieve real-time calculation, the recurrent neural network is adopted for feature analysis, model training is completed on the basis of gradient descent, interpolation calculation of the temperature field in the channel dimension is finally achieved, the resolution of channel dimension data can be effectively improved, then the resolution of a temperature field image is improved, and the method is suitable for large-scale popularization and application. And subsequent thrust chamber temperature field modeling and analysis based on the whole pre-cooling process are facilitated.

The invention discloses a method for designing a mode conversion scheme of a combined propulsion device, which comprises the following steps of: acquiring a Mach number interval of the combined propulsion device converted from a first mode to a second mode, and selecting a plurality of representative flight state points; aiming at each representative flight state point, performing discrete scanning on a mass flow range of a primary propulsion unit and a fuel-air ratio range of a secondary fuel to form a plurality of candidate working condition combinations; flow and combustion characteristics in a thrust chamber are calculated for each candidate working condition combination, and a target working condition combination capable of achieving stable combustion and meeting the thrust requirement is screened; and obtaining an optimal target working condition combination under each representative flight state point based on enumeration search, and generating a continuously enforceable mode conversion scheduling path. The method is applied to the field of combined propulsion, and efficient, stable and safe operation of the mode conversion process can be effectively achieved.

The invention relates to a turbojet engine afterbody comprising:-an exhaust casing (9) carrying a manifold plate (11) surrounding a primary air flow and surrounded by a secondary air flow; -an afterburner (12) comprising a casing (13) and a bushing (16) delimiting together with the casing (13) an annular space (17) extending along the axis (AX), the diameter of the bushing (16) being greater than the diameter of the converging plate (11); -a fixed shroud (22) extending from the bus plate (11) to the inner liner (16), the fixed shroud (22) comprising a window (27); -a moving cover (23) extending along the fixed cover (22) and comprising a window (28), the cover (23) being rotatably movable about an axis (AX) between an open position, in which the window (27) in the fixed cover is open, and a closed position, in which the window (27) in the fixed cover (22) is closed by the moving cover (23).

Additively manufactured thrust chambers and thrust chambers with integral fluid manifolds, and hybrid additive manufacturing methods for their production, are provided. Hybrid additive manufacturing techniques may combine a variety of processes including, WAAM, PBF, cold spray and DED, for example, to produce objects with variant dimensional requirements, i.e., large overall size and small features. Hybrid additive manufacturing may be defined as provide various process layers within any manufactured object. These process layers in turn allow for the introduction of variable feature and size distribution throughout the manufactured object. Hybrid process layers according to aspects may also allow the use of a variety of materials or may use a single material across the various process layers.

The utility model provides a variable-cross-section special-shaped gas exhaust pipe, a liquid rocket engine and a liquid rocket. The variable-cross-section special-shaped gas exhaust pipe comprises an inlet flange arranged at an inlet, an outlet flange arranged at an outlet and a barrel section arranged between the inlet flange and the outlet flange. The inlet and the outlet are eccentrically arranged; the inlet flange is connected with a turbine, and the outlet flange is connected with the thrust chamber gas collecting ring; high-temperature fuel gas enters from the inlet, flows out from the outlet, enters the thrust chamber gas collection ring and then is introduced into the spray pipe extension section. The exhaust pipe is connected between the turbine and the thrust chamber gas collecting ring, the eccentric layout design is adopted, and the problems that in the structural layout of an existing liquid rocket engine, the center distance between a turbine pump and a thrust chamber is large, the radial envelope of the whole machine is large, and the distance between a swing hose and the swing center is large are solved.

The invention relates to a thermodynamic component combined test, and provides a thermodynamic component combined test system and a test method in order to solve the problems that an afterburning cycle engine gas generator is large in propellant mass flow deviation, large in combustion product pollution, high in hot test run cost and the like. The test system comprises a fuel gas generator, a gas venturi tube and a thrust chamber which are coaxially connected; a first component access port is formed in the top of the fuel gas generator, a second component access port is formed in the side face of the fuel gas generator, the first component access port and the second component access port are used for accessing a first component propellant and a second component propellant respectively, and a plurality of fuel gas performance monitoring ports are formed in the side face of the gas venturi tube; a second component regeneration cooling access port is formed in the side surface of the outlet end of the thrust chamber and is used for regenerating and cooling a second component propellant and then accessing the second component propellant into the thrust chamber; the test method based on the test system can effectively improve the efficient, reliable and environment-friendly hot test run test between the fuel gas generator and the thrust chamber.

This application belongs to the field of aero-engine technology, and specifically relates to a double-layer structure afterburner heat insulation plate. The heat insulation plate is installed between the inner cone and the rear bearing cavity, isolating the rear bearing cavity from the inner cone's internal cavity. The heat insulation plate includes an inner layer, an outer layer, a skin, and a mounting edge. The mounting edge is located at the circumferential outer edge of the inner layer and has multiple circumferentially distributed through holes. The mounting edge is installed on the mounting edge of the bottom surface of the inner cone through these through holes. The outer layer is attached to the surface of the inner layer near the inner cone. The inner layer includes a spherical surface protruding into the inner cone's internal cavity. The afterburner heat insulation plate has a double-layer structure, and the cavity height is maintained by fixed supports and rivets. Heat insulation cotton is added to the double-layer cavity to enhance the heat insulation effect of the heat insulation plate.

The application discloses a rocket nozzle and a manufacturing method thereof, comprising a thrust chamber core, a split pressure-bearing shell, an upper sealing cover plate and a lower fastening ring. The application adopts a stainless steel forged hot extrusion integrated thrust chamber core and at least two split shell cold assembly structures, which can adapt to the industrial demand of commercial aerospace rapid iteration, high-frequency test and batch production, with low equipment investment and short processing cycle. The thrust chamber core is a thin-walled Laval tube structure with multiple cooling ribs, which not only reduces the structural dead weight, but also shortens the heat exchange interval between the gas side and the cooling flow channel, greatly improves the heat exchange efficiency, and the cooling ribs can improve the structural rigidity and deformation resistance of the thrust chamber core, so that the rocket nozzle can withstand higher chamber pressure and higher heat flow. The upper sealing cover plate and the lower fastening ring are respectively limited and locked at the top and bottom of the rocket nozzle, which further enhances the structural rigidity of the rocket nozzle, so that the rocket nozzle can further improve the chamber temperature and the chamber pressure, and the working efficiency is higher.

This invention discloses a pneumatic concave cavity flame stabilization device for an afterburner, comprising a central cylinder, an annular combustion chamber shell, a diffuser, an afterburner shell, a pneumatic concave cavity pipe, and a central cylinder connecting rod. The diffuser fixes the annular combustion chamber shell and the afterburner shell together. The central cylinder is fixed to the center of the annular combustion chamber shell via the central cylinder connecting rod. The pneumatic concave cavity pipe extends through the central cylinder into the afterburner shell. The vertical jet from the pneumatic concave cavity exhaust port of this invention can enhance the recirculation zone and stably maintain a low-speed recirculation vortex under different operating conditions, stabilizing the combustion operating range and reducing the risk of flameout. The vertical jet from the pneumatic concave cavity exhaust port of this invention can entrain fuel and air into the cavity, prolonging the fuel residence time and fundamentally improving combustion efficiency. The vertical jet from the pneumatic concave cavity exhaust port of this invention can fill the low-pressure zone at the rear edge of the cavity, reducing resistance, significantly reducing total pressure loss, and improving engine thrust efficiency.

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 application discloses a combined propulsion device mode conversion scheme design method, comprising the following steps: obtaining a Mach number interval of a combined propulsion device when the combined propulsion device is converted from a first mode to a second mode, and selecting a plurality of representative flight state points; for each representative flight state point, respectively performing discrete scanning on a primary propulsion unit mass flow range and a secondary fuel fuel-air ratio range, to form a plurality of candidate working condition combinations; for each candidate working condition combination, calculating flow and combustion characteristics in a thrust chamber, and screening a target working condition combination which can stably combust and meet a thrust demand; obtaining an optimal target working condition combination under each representative flight state point based on enumeration search, and generating a continuous and implementable mode conversion scheduling path. The application is applied to the combined propulsion field, and can effectively realize efficient, stable and safe operation of a mode conversion process.

The invention discloses an overturning and transporting device for a thrust chamber of a liquid rocket engine. The overturning and transporting device comprises a base, an overturning frame, a thrust chamber fixing device, a hydraulic cylinder and a hydraulic control system. The turnover frame, the hydraulic cylinder and the hydraulic control system are mounted on the base; the thrust chamber fixing device is mounted on the roll-over stand; the hydraulic control system is used for driving the hydraulic cylinder to work; the hydraulic cylinder is used for driving the turnover frame to turn over by 0-90 degrees around the rotating shaft; and the thrust chamber fixing device is used for fixing thrust chambers with different sizes. The liquid rocket engine thrust chamber overturning transportation device has the advantages of being easy and convenient to operate, safe, efficient, high in adaptability and the like, and the problems that a traditional overturning transportation mode is high in safety risk, low in efficiency and the like are effectively solved.

The invention discloses a double-thrust-chamber pump rear-swinging liquid rocket engine, relates to the technical field of rocket engines, and provides a technical scheme capable of effectively making up for the defects of a traditional single-turbine pump and single-thrust-chamber structural layout in a high-thrust or heavy-duty liquid rocket engine. The invention discloses a double-thrust-chamber pump rear swing liquid rocket engine which comprises a rack, two fuel gas swing devices, two thrust chambers and a fuel gas assembly. The fuel gas assembly communicates with the corresponding thrust chambers through the two gas swing devices and is used for providing fuel for the two thrust chambers correspondingly. The first ends of the two fuel gas swing devices are fixedly arranged on the two sides of the rack correspondingly, the second ends of the two fuel gas swing devices are fixedly connected with the corresponding thrust chambers correspondingly, and the two fuel gas swing devices and the two thrust chambers are symmetrically distributed along the center line of the rack. The fuel gas assembly is fixedly arranged in the middle area of the rack.

This invention relates to a gas generator circulating liquid rocket engine split-flow drive system and method. The drive system includes a gas generator, a fuel pump, a fuel pump turbine, an oxidizer pump, an oxidizer pump turbine, and a thrust chamber. The fuel pump turbine is coaxially and fixedly connected to the fuel pump, and the oxidizer pump turbine is coaxially and fixedly connected to the oxidizer pump. The gas generator is connected to both the fuel pump turbine and the oxidizer pump turbine via pipelines. The fuel pump is connected to a first fuel pipeline and a second fuel pipeline. The first fuel pipeline is connected to and communicates with one of the gas generator and the thrust chamber, and the second fuel pipeline is connected to and communicates with the other of the thrust chamber and the gas generator. The oxidizer pump is connected to a first oxidizer pipeline and a second oxidizer pipeline. The first oxidizer pipeline is connected to and communicates with one of the gas generator and the thrust chamber, and the second oxidizer pipeline is connected to and communicates with the other of the thrust chamber and the gas generator.

The afterburner comprises an outer casing, a center body assembly and an oil way assembly, the center body assembly and the oil way assembly are arranged in the outer casing, the center body assembly comprises an inner casing, a center cone located in the inner casing, a rectification supporting plate located between the inner casing and the center cone and an outer concave cavity, and the outer concave cavity forms a protrusion in the direction from the inner casing to the outer casing. The rectifying supporting plate comprises a plurality of rectifying blades, grooves are formed in the wall faces of one sides of the rectifying blades, the grooves of the rectifying blades extend from the center cone to the outer concave cavity, the grooves of the rectifying blades are located in the middles of the rectifying blades in the axial length direction and correspond to the outer concave cavity in the radial direction, and the oil spraying holes of the oil way assembly are located in the rectifying blades. The grooves are formed in the rectifying blades and matched with the outer concave cavities to jointly complete ignition, flame transfer and flame stabilization of the afterburner, the combustion position is organized in advance, the whole afterburner can have the shorter length-diameter ratio and the lighter structural weight on the premise that thrust is not sacrificed, and the thrust-weight ratio is increased.

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 application relates to a posture-orbit control thrust chamber body structure with coating protection and a preparation method thereof, and solves the technical problem that the preparation of the coating of a high-thrust high-performance posture-orbit control thrust chamber body is limited by the size of an ion sputtering furnace and the ion sputtering coating with better high-temperature resistance and oxidation resistance cannot be used. The posture-orbit control thrust chamber body structure with coating protection comprises a short body part and an extension section which are sequentially connected; an ion sputtering coating is arranged on the outer wall of the short body part; and a slurry sintering coating is arranged on the outer wall of the extension section, so that the ion sputtering coating can have good high-temperature resistance and oxidation resistance, and the requirement of segmented protection of different regions of the high-thrust high-performance posture-orbit control thrust chamber body can be met. The coating preparation method of the posture-orbit control thrust chamber avoids the problem that the base body of the posture-orbit control thrust chamber is not protected by the coating.

The application relates to a thrust chamber of a jet propulsion device and an adaptive flame stabilizer thereof. The adaptive flame stabilizer comprises two side plates connected in rotation, an oil injection rod and a linkage mechanism. The oil injection rod is arranged between the two side plates, and a first nozzle is arranged on the oil injection rod. The linkage mechanism comprises an extension assembly and two linkage assemblies. One end of the extension assembly is communicated with the first nozzle, and the other end of the extension assembly is rotatably connected with the two linkage assemblies respectively. The ends of the two linkage assemblies away from the extension assembly are rotatably connected with the two side plates respectively. The adaptive flame stabilizer has a thrust state and a non-thrust state. In the thrust state, the oil injection rod sprays oil to the first nozzle, and the extension assembly is elongated to drive the two side plates to rotate to the side away from each other through the two linkage assemblies respectively. In the non-thrust state, the extension assembly is retracted to drive the two side plates to rotate to the side close to each other through the two linkage assemblies respectively.

This application discloses a CuCrNbZr alloy material, its preparation method, and its applications. The CuCrNbZr alloy material, by mass percentage, comprises: Cr: 2.2%~3.5%, Nb: 1.8%~3.0%, Zr: 0%~0.15%, with the balance being Cu. This CuCrNbZr alloy material is prepared through vacuum atomization powdering, hot isostatic pressing, and heat preservation cooling, and can be applied to the thrust chamber and nozzle of liquid rocket engines. Through the synergistic effect of specific component ratios and preparation processes, this application achieves a yield strength ≥150MPa, tensile strength ≥300MPa, and strength retention rate ≥85% after high-temperature brazing at 935℃~975℃. This solves the problems of rapid strength drop and insufficient structural reliability of traditional copper alloys after brazing, meeting the high-temperature and high-pressure requirements of hot-end components in liquid rocket engines.

This application relates to a device for reducing the servo load torque of a reusable rocket's return stage engine, comprising a central engine and multiple skirts. These skirts are foldable and uniformly arranged circumferentially on the outer wall of the central engine's thrust chamber. This application offers the following anticipated technical advantages: the device can significantly reduce the aerodynamic load torque of the central engine's servo system during the return phase, greatly reducing the design complexity of the central engine servo system and lowering the development cost of the reusable rocket; the skirts are folded during the central engine's ascent and return phases, fitting snugly against the outer wall of the central engine's thrust chamber, without affecting the central engine's oscillation operation; the skirts are controlled by linear servos, resulting in a simple structure that can be freely folded and unfolded.

This invention discloses a combined propulsion system and multi-modal operation method based on liquid zero-carbon fuel. The combined propulsion system includes a first propulsion assembly, a second propulsion assembly, a pre-combustion chamber, a liquid ammonia storage tank, and an oxidizer storage tank. The first propulsion assembly includes an air intake channel, an isolation section, a first thrust chamber, and a first nozzle connected in sequence. The second propulsion assembly includes a second thrust chamber and a second nozzle connected in sequence. The liquid ammonia storage tank is connected to the first thrust chamber and the pre-combustion chamber via a fuel supply channel assembly. The oxidizer storage tank is connected to the second thrust chamber and the pre-combustion chamber via an oxidizer supply channel assembly. The pre-combustion chamber is connected to the first thrust chamber and the second thrust chamber via a gas supply channel assembly. This invention is applied in the field of combined propulsion and can intelligently adjust the cooling path and cracking strategy of liquid ammonia according to different operating modes, achieving efficient thermal management and stable combustion of a single liquid ammonia fuel in a wide-speed range of flight.

The invention discloses a high-thrust extrusion type rocket manufactured based on a mature industrial system. The high-thrust extrusion type rocket comprises a fuel storage tank, an oxidizing agent storage tank, a liquid nitrogen storage tank, an exhaust pipe, an oxidizing agent injector, a fuel injector and a thrust chamber. The nitrogen turbine pump is only used for pressurizing inert gas, an oxidizing agent and fuel are directly extruded into the thrust chamber, the risk of explosion caused by rapid oxidation or leakage is avoided, meanwhile, the turbine is driven by high-temperature fuel gas extracted from the thrust chamber, a combustor is not needed, the structure of the turbine pump is simplified, and the cost is reduced. In addition, the nitrogen turbine pump is driven by high-temperature gas and used for pressurizing nitrogen, batch production or direct purchase of the nitrogen turbine pump can be achieved depending on an existing mature industrial system, and the manufacturing cost of the nitrogen turbine pump is remarkably reduced; in addition, under the pressurization effect of the nitrogen turbine pump, the propellant is effectively extruded to enter the thrust chamber in a large-flow mode, a high-pressure gas cylinder is not needed, the three storage boxes do not need to be provided with high-pressure thick walls, and the overall structure manufacturing cost of the rocket is remarkably reduced.

The application discloses a surface treatment device for an SLM engine thrust chamber and relates to the technical field of additive manufacturing. The surface treatment device comprises a polishing cabin and a control system for controlling polishing treatment in the polishing cabin. The polishing cabin is internally provided with an inner cavity annular surface treatment mechanism, a single-grinding-head surface treatment mechanism, an environment control mechanism, a base component, a pneumatic adjusting mechanism, a hydraulic control mechanism and an SLM forming engine thrust chamber workpiece. The application overcomes the shortcomings of the prior art, adopts mechanical processing to improve the efficiency and consistency of production, and adopts a digital control system to realize controllable adjustment of the roughness of the inner and outer surfaces of the SLM engine thrust chamber.

A nozzle for a vehicle is disclosed. The nozzle includes a plurality of thrust chambers spaced relative to one another. Each thrust chamber has a throat and a nozzle portion extending aft of the throat. The nozzle portion has an exit dimension Dexit at an aft end. Each thrust chamber is spaced away from each adjacent thrust chamber by a respective spacing distance Dspace, such that Dspace≥M*Dexit, where M≥1.

The application belongs to the technical field of aero-engines, and particularly relates to a layout structure of a radial stabilizer of a thrust chamber, a casing, a converging ring, an inner cone body and a plurality of support plates distributed circumferentially between the converging ring and the inner cone body; wherein the tail edge of the support plate is a long radial stabilizer with a gradually increasing circumferential width along the airflow direction, the converging ring is provided with a short support plate at the position between adjacent long radial stabilizers, the fixed end of the short support plate is fixed on the inner surface of the converging ring, the free end extends radially inward by a certain length, the tail edge of the short support plate is a short radial stabilizer with a gradually increasing circumferential width along the airflow direction, the fixed end of the short radial stabilizer is fixed on the inner surface of the converging ring, and the free end extends radially inward by a certain length; the tail end of the long radial stabilizer and the tail end of the short radial stabilizer have the same axial position, the application increases the backflow area of the tail end of the stabilizer, and is beneficial to the radial propagation and combustion of the flame.

The invention discloses a support plate stabilizer structure for enhancing fuel oil atomization and suitable for an afterburner, which is used for improving the fuel oil atomization effect and the combustion stability, and comprises two support plate outer edge frames, a support plate tail edge frame, a fuel oil injection rod and the like, and a support plate mixing channel is formed between the two support plate outer edge frames; and a plurality of cooling air cavities are formed in the main body part so as to prevent key parts from being ablated in a high-temperature environment. First fuel oil impacts the fan-shaped channels through the oil injection holes to form an oil film, is further atomized under the shearing action of airflow and is mixed with air, and premixed gas is formed and then enters the concave cavity to be combusted. According to the design, flames in the concave cavity can be kept stable under the extreme conditions of low temperature, low pressure and the like, so that the ignition and flameout boundary of the combustion chamber is widened, meanwhile, high-temperature fuel gas in the concave cavity can heat transverse jet flow of second fuel oil, and the evaporation and mixing efficiency of the second fuel oil is improved. By means of the integrated design, the number of components is reduced, the flow resistance in the non-stress-application state is reduced, and meanwhile the thrust-weight ratio is increased.

The gas propulsion thrust device comprises a cone-shaped propulsion element with a rigid concave internal surface and a second convex external surface. The propulsion element is submerged in gas and aligned with a high-frequency linear actuator, which causes its reciprocal motion along the longitudinal axis, generating thrust. The device includes a thrust chamber that supports the actuator and surrounds the propulsion element, maintaining a consistent gap between the propulsion element and the chamber to direct gas from the second side to the first side. The chamber tapers around the propulsion element, and a gas directing cap is configured to direct gas around the second side towards the first side. Propulsion element is rigid and its shape remains unchanged during the operation. The reciprocal motion creates a gas pressure differential across the propulsion element, generating thrust by propelling gas away from the propulsion element in the opposite direction of propulsion. The device offers efficient and controlled thrust generation.

The application discloses a kind of shunt thrust control coaxial double-paddle aircraft and flight control mode, it is related to aircraft technical field.The application includes: coaxial double-paddle unit and thrust control unit, thrust control unit, the thrust control unit includes thrust chamber and multiple groups of flow regulating mechanism, the thrust chamber is placed below the lower rotor and is connected with the mounting frame, it is circumferentially spaced apart and is provided with multiple through-flow cavities along the axial direction of the upper rotor, multiple groups of flow regulating mechanism are correspondingly installed in multiple the through-flow cavities, to adjust the flight attitude of aircraft by changing the flow area corresponding to the through-flow cavity to adjust the resistance size when airflow passes through each the through-flow cavity to adjust the thrust distribution in thrust chamber.The application combines the advantages of coaxial double-paddle aircraft and multicopter, so that the aircraft can obtain larger paddle radius and lift action area, and has good controllability, maneuverability and stability.