53 results about "Turbojet" patented technology

A method for manufacturing an alveolar core structure includes at least one cell including a secondary duct having a first end defining a sound wave inlet, and an opposite second end, the secondary duct comprising a sound wave outlet. The method also includes a fastening step in which adhesive tapes transverse to the longitudinal direction of said first plate are applied on a first longitudinal plate. The secondary duct in the form of a flattened element is fastened, on the first plate, by gluing at its sound wave inlet. A second plate is applied. A step of deploying the first and second plates so as to form the peripheral wall of the cells and so that the flattened element is deployed.

The invention provides an aero-engine thermal jet flow FT-IR spectrum enhancement classification method, and the method carries out the infrared spectrum detection of six different types of aero-engine tail jet thermal jet flows including two types of turbojet and turbofan, serves as the basis of engine recognition and data source input, and is more scientific in classification. Meanwhile, aiming at the problem that the CNN method has overfitting on the existing spectral data set, the invention provides an FTIR-SpectraGAN method, new data is generated through the existing data, the classification capability of the network in various aspects is improved, and higher classification accuracy is achieved.

The unmanned wheeled vehicle, the control method and the control system thereof can assist the unmanned wheeled vehicle to overcome the obstacle and get rid of the predicament when the unmanned wheeled vehicle is trapped by the obstacle, thereby improving the adaptability of the unmanned wheeled vehicle to the complex terrain and improving the maneuverability of the unmanned wheeled vehicle. The unmanned wheeled vehicle comprises an unmanned vehicle main body structure, an auxiliary power system connected with the unmanned vehicle main body structure and configured to assist the unmanned vehicle main body structure to overcome the obstacle, wherein the auxiliary power system comprises a turbojet system, an unmanned vehicle state detection system configured to detect the state of the unmanned vehicle main body structure, and a control system electrically connected with the unmanned vehicle state detection system and the auxiliary power system and configured to control the auxiliary power system according to the state detection result of the unmanned vehicle state detection system.

This invention relates to the field of underground coal mining technology and discloses an explosion-proof high-energy dynamic current power generation device for underground coal mines based on non-contact turbojet technology. The device includes a pipeline, with a support frame fixedly connected to the inner ring of the pipeline. A transmission rod is rotatably connected to the inner ring of the support frame, with one end fixedly connected to a housing. A main rotor is mounted on the other end of the transmission rod. A support plate is fixedly connected to one side of the inner ring of the pipeline, and a generator is fixedly connected to one side of the outer wall of the support plate. An active bevel gear is connected and fixedly attached to the input end of the generator. Based on wind power meter readings, an electric push rod is used to disengage the corresponding main rotor from the slave rotor. Compressed air inside the pipeline then drives the housing to rotate, generating electricity. If the wind power meter reading is insufficient, a branch pipe is opened to output water from a power waterwheel, which then drives the waterwheel to rotate, generating electricity. This allows for flexible switching based on actual conditions.

The invention belongs to the field of afterburners of aero-engines, and particularly relates to an afterburner oil supply device. The oil supply device comprises an oil supply main body, a fuel oil flow channel is formed in the oil supply main body and comprises two or more fuel oil flow channel sections which are sequentially communicated, an inlet of one fuel oil flow channel section is communicated with an oil supply pipeline, a plurality of oil spraying holes are formed in the oil supply main body, and the oil spraying holes are communicated with the fuel oil flow channel sections. The fuel injection holes are communicated with the corresponding fuel runner sections; the fuel oil supply assemblies are in one-to-one correspondence with the fuel oil flow channel sections, and all the fuel oil supply assemblies are arranged in the corresponding fuel oil flow channel sections respectively. The radial partition oil supply and combustion requirements of the afterburner can be met, and better afterburner oil-gas mixing and combustion effects are provided; and the structure design is more flexible, and finer fuel injection control and a more accurate oil-gas matching effect can be achieved.

The invention discloses a water-air dual-purpose engine. The engine takes air or water as a combustion medium; the engine includes: a medium introduction part for introducing a combustion medium into the engine; the trapped vortex combustion part is used for accommodating the fuel and the combustion medium under the condition that the combustion medium is air, so that the fuel is combusted; the metal powder combustion part is used for accommodating fuel and a combustion medium under the condition that the combustion medium is water, so that the fuel is combusted; and the substance spraying part is used for spraying substances generated by combustion out of the engine so as to generate traction force. According to the water-air dual-purpose engine, the two different propelling modes of turbojet and metal powder water stamping are integrated, the two propelling modes are compatible at the same time according to different environments and working conditions, and the working states are synchronously switched during environment switching so as to adapt to combustion media corresponding to the different environments; therefore, the engine can provide power for the aircraft in different propulsion modes in the air environment and the water environment respectively, and the stability of fuel combustion is kept.

The application discloses a hyperplane climbing track group intelligent optimization method, which comprises the following steps: constructing a hyperplane climbing section dynamics equation, and obtaining to-be-optimized parameters in the hyperplane climbing section dynamics equation; designing a hierarchical cost function considering multiple constraints and turbojet, subsonic ramjet and supersonic ramjet three power mode switching Mach numbers; and optimizing the cost function through a chicken swarm algorithm, thereby completing the hyperplane climbing track group optimization. The hyperplane climbing track parameter optimization is performed through the improved chicken swarm algorithm, the suitable power mode switching Mach number is selected while the complex constraints of the hyperplane climbing section are met, the climbing scheme capable of significantly shortening the climbing time and total fuel consumption is obtained, and the problem that the track is difficult to optimize under the hyperplane cross-power mode and long-time window climbing condition is solved.

1. Name of the product in this design: Thrust Test Platform for Turbojet and EDF Engines. 2. Purpose of this design: This design is used to measure the thrust ratio of the EDF engine thrust test platform. 3. The key design feature of this product is its shape. 4. The image or photograph that best illustrates the design's key points: 3D view 1.

The application relates to the field of general structure design of an aero-engine, and discloses an aero-engine duct airflow distribution mechanism and a control method thereof. The bleed air structure and the flow path control structure are integrated in the rotating casing and the support plate structure of the inner and outer ducts. In the turbofan mode, the leading edge and the trailing edge of the corresponding support plate of the outer duct and the inner duct are matched, so that the duct flow is smooth. In the turbojet mode, the inner and outer duct flow passages are blocked, and the outer duct airflow is guided into the inner duct. The application can support the demand that the variable flow path aero-engine works in the turbofan engine mode at low flight speed and works in the turbojet mode at high flight speed, widen the flight speed domain of the conventional turbofan engine, realize the consideration of high economy at low speed and large thrust at high speed, and effectively reduce the structural complexity of the engine.

The invention discloses a hybrid electric engine using electric supercharging and indirect cooling. The engine is based on a turbojet or turbofan core engine. An electric supercharging indirect cooling device and a turbine power generation device are additionally arranged on a core engine and are installed at the front end and the rear end of the engine respectively. The electric supercharging intercooling device comprises an electric supercharging channel with a contracted flow channel, an electric driving fan is used in the channel for supercharging, the fan is a wide-pressure-ratio working fan, a motor for driving the fan is located in the center of the flow channel at the rear end of the fan, the flow channel is narrowed after the motor passes through the electric supercharging fan, and a compact intercooler is installed on the rear portion of the flow channel. A flange is arranged at the casing behind the intercooler and is connected with a shell of the core engine; the rear-end turbine power generation device is connected with an outlet of the core engine through a flange, the rear-end device is provided with an adjusting device and divides a duct into an inner duct and an outer duct, the inner duct is provided with a turbine, the turbine is connected with a power generator through a force transmission shaft perpendicular to the direction of a flow channel after passing through a direction changing device, and the power generator is installed outside the flow channel of the device.

The invention relates to a turbomachine turbine comprising a casing (12) carrying a distributor (13) extending circumferentially around an axial direction, in which: – this distributor is formed of several bladed sectors (22, 23, 24) each comprising a radially external platform (26, 27, 28) and a radially internal platform with several blades (14) extending radially from the internal platform to the external platform (26, 27, 28); – the casing (12) comprises an internal groove (29) of revolution, with several distinct ring sectors (32, 33, 34) each engaged in this internal groove (29) and protruding radially towards the inside of the casing (12); – each ring sector (32, 33, 34) axially blocks at least one bladed sector (22, 23, 24). Figure for abstract: Figure 5

The invention provides an aviation turbojet and turbofan engine corrosion resistance test device and test method. The device comprises an air inlet unit, a temperature adjusting unit, a humidity adjusting unit, a salt mist mixing adjusting unit, a dynamic salt mist sampling unit, a salt mist protection cabin, an air adjusting unit, an exhaust unit and a measurement and control unit. In the engine operation stage, the exhaust injection effect of a test engine is used as a system air inlet main driving power source, and the air supply flow stability of the system in different operation states of the engine is realized by combining the adjustment of an exhaust fan; different flow path designs are carried out according to test requirements of different stages of an engine corrosion resistance test; and the risk of system control failure caused by the fact that a sensor in the protection cabin is low in reliability due to the long-term salt mist airflow effect is reduced while it is guaranteed that the environment in the protection cabin meets the test requirement.

The application discloses an aero-engine hot-end component full-life-cycle management method and system, and belongs to the technical field of aero-engines.The system constructs a full-life-cycle closed loop system of material gene optimization-precision forming process-service health monitoring-life prediction and active repair, can realize service life improvement of more than 200% of the aero-engine hot-end component, and improves the production yield rate from the conventional 60% in the industry to more than 90%.The application fills the core short board of short life, poor reliability and low yield rate of the aero-engine hot-end component in China, breaks through the core technical bottleneck of the "industrial crown pearl" of the aero-engine, and is 100% domestically produced and controllable in material and process, and can be widely applied to scenes such as army armed helicopter turboshaft engines, military unmanned aircraft turbojet / turbofan engines and the like.

This invention discloses a method for analyzing the transient characteristics of multi-ducted engine casings under complex loads, belonging to the field of aero-engine technology. This method analyzes the structure and operation of a multi-ducted variable cycle engine, classifying it into three turbofan modes and one turbojet mode. Based on an aerodynamic thermodynamic configuration model, component-level modeling is performed using MATLAB software, generating component-level modeling data. A common set of working equations is established using the modeling data, and the solution is obtained through iterative solving using the Broyden quasi-Newton method. The ultimate aerodynamic loads of the casing system under typical operating conditions are calculated. Combined with the engine flight envelope, 12 typical altitude and Mach number states are selected to solve for the ultimate operating points and analyze the transient load characteristics. This invention can accurately capture the transient response characteristics of multi-ducted engine casing systems under complex load conditions, improving analysis accuracy and efficiency, providing a theoretical basis for the design optimization and reliability assessment of casing systems, and has significant engineering value.

1. Name of the product in this design: Central Controller. 2. Purpose of this design: To control the operation of a turbojet snowplow and to display real-time dynamic information feedback from the turbojet snowplow. 3. The key design feature of this product is its shape. 4. The image or photograph that best illustrates the design's key points: a 3D model.

The application provides a nuclear energy electric supercharged jet propulsion system and belongs to the field of aviation power. The nuclear power propulsion system of the traditional turbine engine type solves the problems that the working processes of a compressor and a turbine influence each other, the matching of the compressor and the turbine is relatively difficult, the performance of the propulsion system under variable working conditions is poor, and the arrangement of a heat exchanger in a closed cycle nuclear power system is relatively complex and the flow resistance of the heat exchanger is relatively large. The nuclear energy electric supercharged jet propulsion system comprises a regenerator, a cold end inlet connected with an outlet of a helium-xenon compressor, a cold end outlet connected with a cold end inlet of a heater, a hot end outlet of a first heat exchanger connected with a hot end inlet of a fuel heat exchanger, a second heat exchanger, a hot end inlet of the second heat exchanger connected with a circulating working medium outlet of a reactor system, a fuel tank, a discharge port of the fuel tank connected with the cold end inlet of the fuel heat exchanger through a fuel valve, and an engine shell, wherein air from the compressor outlet flows through the first heat exchanger and the second heat exchanger in sequence. The nuclear energy electric supercharged jet propulsion system is mainly used for the propulsion system of aviation equipment.

The invention discloses an assistant platform of a turbojet fixed-wing aircraft, and relates to the related technical field of aircrafts. Comprising a tilting mechanism, a supporting mechanism, a moving mechanism, a rotating shaft, a plurality of guide shaft fixing supports, a fixed pulley block, a manual winch, a steel wire rope, a gear and a chain. The manual winch is shaken, the chain drives the gear to rotate, the inclination mechanism is turned over upwards, and then the inclination angle of the turbojet fixed-wing aircraft can be changed. When the gravity center of the tilting mechanism shifts due to tilting, mechanism resetting is achieved through the synergistic effect of traction force of the steel wire rope and gravitational potential energy, then angle adjustment of the tilting mechanism from 0 degree to 90 degrees is achieved, controllable ignition of aircraft turbojet within the inclination angle range from 0 degree to 90 degrees can be guaranteed, and the ignition process is stable and reliable. A high-credibility experimental environment is constructed for functional verification of the turbojet aircraft, a research platform is provided for optimization and iteration of the turbojet fixed-wing aircraft, and research and development test efficiency and safety are effectively improved.

The present application belongs to the field of optimal control technology, and particularly relates to a combined power boost reusable spacecraft climbing segment trajectory optimization method, aiming at the problem that in the existing trajectory optimization of the climbing segment of the reusable spacecraft, the processing of the thrust is idealized, and the speed improvement effect and the height improvement effect are not matched; on the basis of the classical Gauss pseudospectral method, in order to improve the speed of solving the non-smooth problem, in the polynomial degree and collocation point correction, the adaptive decision method is introduced based on the curvature of the control variable on the time period, the traditional Gauss pseudospectral method is combined with the finite element method, and the method of increasing the interpolation polynomial degree or increasing the time interval is used for correction. Under the premise that the turbojet and ram work together, the distribution mode of the action time and the thrust size of the two is proposed, and the distribution scheme of the combined thrust is given.

A turbojet manned aircraft based on vector control belongs to the field of aircraft. In order to solve the problems of complex attitude adjustment and large amount of calculation in existing manned aircraft. The two vector power units in the present invention are arranged along the x-axis direction and symmetrically installed on the left and right sides of the fuselage, and provide thrust for the navigation and attitude adjustment of the aircraft; the man-machine fixing unit and the fuel tank are arranged along the y-axis direction and are respectively installed on the front and rear sides of the fuselage, and the man-machine fixing unit is fixed on the operator's body to achieve the fixation of the aircraft and the operator; the fuel tank is respectively connected to the two vector power units and provides endurance fuel for the two vector power units; the main control unit is installed on the fuselage 1 and is respectively electrically connected to the two vector power units to achieve the control of the thrust vector of each vector power unit, thereby achieving the control of the thrust vector of the aircraft. The present invention is mainly used for personnel transportation.

A simple two-dimensional vector nozzle relates to the field of turbojet or turbofan engine power control, specifically to a device for controlling the thrust direction of the engine. The mechanism includes components such as a support, a cylindrical sleeve, a large push rod, a conical sleeve, and a small push rod. The support can be fixed to the engine or other structural parts around the engine to fix the entire mechanism. The cylindrical sleeve rotates relative to the support via large ball bearings on both sides. The large push rod is welded to the cylindrical sleeve. The conical sleeve is connected to the cylindrical sleeve via a small ball bearing through an axial fit and rotates relative to the cylindrical sleeve. The small push rod is welded to the wall of the conical sleeve. By driving the large push rod and the small push rod to move simultaneously, the cylindrical sleeve and the conical sleeve are driven to rotate simultaneously, achieving compound motion of the vector nozzle in two degrees of freedom, front-to-back and left-to-right, and realizing thrust direction control.

The utility model relates to the technical field of annealing furnaces, and discloses a multi-blade fan-shaped turbojet type circulating air device which comprises a volute, and a fan is mounted on the volute; a left air cavity is fixedly connected to the left side of the volute, a right air cavity is fixedly connected to the right side of the volute, a plurality of left air nozzles are arranged on the right side of the left air cavity in an inclined mode, and a plurality of right air nozzles are arranged on the left side of the right air cavity in an inclined mode. An air guide plate is fixedly connected to the interior of the volute, an air suction flow guide cover communicated with the volute is fixedly connected to the bottom of the volute, and a flow guide disc is fixedly connected to the interior of the air suction flow guide cover. According to the multi-blade fan-shaped turbojet type circulating air device, the left tuyeres and the right tuyeres are arranged in the annular array respectively and form certain included angles with the left air cavity and the right air cavity, so that sprayed air flow is in a large-rotation vortex, and a material coil can be heated by hot air more uniformly.

The utility model relates to the field of bearing devices, and discloses a special bearing for a low-altitude turbojet aircraft, which comprises an outer ring, a filling port is arranged on the side wall of the outer ring, an oil storage tank is arranged inside the outer ring, second grooves are respectively arranged on two sides of the inner side wall of the outer ring, a plurality of lubricating ports are respectively arranged on the inner side walls of the two second grooves in a circumferential array mode, and the lubricating ports are communicated with the oil storage tank. A plurality of rollers are slidably connected to the inner side walls of the two second grooves correspondingly, a sealing plug is slidably connected to the inner side wall of the filling port, an inner ring is arranged on the inner side wall of the outer ring, and first grooves are formed in the two sides of the side wall of the inner ring correspondingly. According to the special bearing for the low-altitude turbojet aircraft, the two rows of parallel pin rollers are arranged in the outer ring, so that higher rotation precision and lower vibration level are achieved, and the oil storage tank is formed in the outer ring, so that material performance reduction or bearing failure caused by overheating is prevented.

The invention relates to the field of aero-engine design, and discloses a distributed hybrid power layout system and a control method, and the system comprises a gas turbine main body, a bypass air entraining casing, a valve, a free turbine, a generator and a ducted fan. The gas turbine engine and the ducted fan are mechanically decoupled through electric energy, the ducted fan replaces a low-pressure component of a traditional gas turbine engine, and an inner culvert and an outer culvert of the gas turbine engine are separated on the structural layout, so that design constraints are reduced, and flexible distributed layout and large-range adjustment of the bypass ratio are achieved; working mode switching of the aircraft is achieved by opening or closing the valve for opening the bypass air entraining casing, the propulsive efficiency can be improved in a large bypass ratio turbofan mode under the cruising working condition, unit thrust can be improved in a turbojet mode under the supersonic speed working condition and other working conditions, and therefore performance improvement in a wide range is achieved.

A light speed adjustable laser interference system applied to a turbojet nozzle shell shape righting tool comprises a Michelson interference system and the shape righting tool, the shape righting tool comprises a bottom plate, a front plate, a rear plate and an upper plate, the front plate, the rear plate and the upper plate form a mounting frame, a first through hole is formed in the upper portion of the mounting frame, an upper pressing shape righting mechanism is assembled in the first through hole, and a second through hole is formed in the lower portion of the mounting frame. The bottom plate is provided with a Michelson interference system, the Michelson interference system comprises a light stroke difference and light velocity fine tuning system, the light stroke difference and light velocity fine tuning system comprises a glass tube, an oil container and a bidirectional pump, the side wall of the glass tube is provided with an oil supply port, the glass tube is vertically arranged on a light path between the beam splitter and the reflective mirror, and the oil supply port is communicated with the bidirectional pump. And a floating transparent sheet with the density smaller than that of the light-transmitting oil is arranged in the glass tube. The optical path difference is adjusted in the mode that the liquid level in the light path difference and light speed fine adjustment system rises and falls, the equal optical path is achieved between the reflection reflective mirror and the transmission reflective mirror, and the positioning, measuring and shape correcting precision is greatly improved.

The invention belongs to the technical field of aero-engine testing, and discloses a magnetic levitation supported wide-range high-precision small aero-engine thrust measuring rack. A movable frame module and a fixed frame module of the thrust measuring rack are both of a horizontally-placed flat plate structure, a small turbine jet aero-engine test piece to be measured is fixed to the upper portion of the movable frame module, and the fixed frame module is fixed to a test bench base. The movable frame module and the fixed frame module are supported through four vertical supporting cylinders arranged at the four corners. A pair of vertical floating magnetic steel assembly and fixed magnetic steel assembly is arranged in each supporting cylinder; and a multi-gear range selection module and a multi-gear in-situ calibration module are also fixed on the upper surface of the fixed frame module. The thrust measuring rack is wide in thrust measuring range, large in measuring range ratio, high in measuring efficiency, high in calibration precision, high in calibration efficiency and good in economical efficiency, can meet the thrust measuring requirements of the wide-range small turbine jet aero-engine, and has engineering practical value.

A long-range heavy-lift turbojet-powered manned vertical take-off and landing aircraft layout, comprising: a riding cockpit, telescopic variable-span wings, landing gears, turbojet thrust vectoring lift engine compartments, and a turbojet thrust engine. The riding cockpit, as a structural body, is located at the center of the aircraft; the bottom of the riding cockpit is connected to the telescopic variable-span wings; four turbojet thrust vectoring lift engine compartments are vertically arranged at the front, rear, left and right corners of the cockpit connected to the telescopic variable-span wings; the turbojet thrust engine and an air intake / exhaust duct thereof are horizontally embedded in the middle area between the wings in the front-back direction; and the landing gears are arranged on two sides of the turbojet thrust engine. The telescopic variable-span wings consider the take-off and landing stability and the cruise economy requirements by means of telescopic change; in addition, by additionally providing end plates at wing tips, the cruise flight induced resistance can be reduced, and the wing tips can be prevented from being scratched. The present invention considers heavy-lift / long-range flight capability, controllability and stability, and achieves pilot-aircraft attitude decoupling; and a pilot can control the aircraft attitude on the basis of body sensing, such that the operation requirement for the pilot is low.

The invention discloses a multi-duct variable-cycle engine maneuvering and aerodynamic load prediction method and system, and relates to the technical field of aero-engines. The method comprises the steps that typical flight actions including climbing, diving and level flight are divided based on flight tasks, a rain flow counting method is adopted to filter an original load spectrum so as to retain a key loading and unloading path, three-parameter Weibull distribution is fitted in combination with a bilinear regression method, and load distribution characteristics of typical task segments are extracted; establishing a flight dynamics equation and a lift-drag characteristic model, and verifying the prediction precision of the normal overload coefficient through Simulink simulation; and the maneuvering and aerodynamic load distribution of the multi-duct variable-cycle engine in a turbofan / turbojet mode is deduced by further combining the airfoil profile parameters of the high-sound-velocity aircraft. According to the method, the accuracy of load prediction can be remarkably improved, a reliable basis is provided for service life evaluation and optimization design of an engine structure, and the method has important engineering value.

The present invention discloses a lobe-type rear duct ejector with modal adjustment and mixing enhancement functions, comprising a variable geometry lobe, a hydraulic actuator, a transmission mechanism, an outer casing, an intermediate casing, and a central cone. The variable geometry lobe is composed of a fixed crest and an adjustable trough, and the lobe sidewalls are a foldable structure that can adjust the flow area of ​​the duct outlet according to the ejection requirements of the outer duct airflow. The sidewall receiving cavity on the outer casing can reduce the limitation of the adjustable trough's outward deflection, helping the variable geometry lobe to further reduce the lower limit of the bypass ratio adjustment, thereby better leveraging the high-speed performance of the near-turbojet mode. During the transition to the turbofan mode, as the folded sidewalls unfold, the leeward side of the sidewalls can induce the jet wake to generate large-scale flow vortices, thereby enhancing mixing and increasing the thrust gain of the mixed jet. The present invention can meet the performance requirements of ejectors of different modes, enabling the rear duct ejector to maintain excellent thrust performance under all operating conditions.

The application belongs to the field of turbojet afterburner, and particularly relates to an afterburner oil supply device. The oil supply device comprises: an oil supply main body, an oil flow channel is formed in the oil supply main body, the oil flow channel comprises two or more oil flow channel sections which are sequentially communicated, an inlet of one of the oil flow channel sections is communicated with an oil supply pipeline, a plurality of oil injection holes are formed in the oil supply main body, and the oil injection holes are communicated with corresponding oil flow channel sections; and oil supply assemblies which are one-to-one corresponding to the oil flow channel sections, each oil supply assembly is arranged in a corresponding oil flow channel section. The application can match the radial partition oil supply and combustion requirements of the afterburner, provide better oil-gas mixing and combustion effects of the afterburner, and has a more flexible structure design, so that more fine fuel injection control and more accurate oil-gas matching effects can be realized.

The invention discloses an axial flow-centrifugal combined type turbojet propulsion power device based on coupling of a coupler and an operation method of the axial flow-centrifugal combined type turbojet propulsion power device, and belongs to the technical field of aviation power, and the device is formed by coaxially connecting a centrifugal type core machine unit body and an axial flow type combined unit body through the coupler. An engine body extends forwards in the axis direction, an axial-flow type combined unit body is additionally arranged in the extending engine body through a coupler, and specifically components such as a fairing assembly, a fuel oil and lubricating oil combined pump, an axial-flow compressor assembly and a built-in generator are formed. The invention further provides a corresponding operation method which comprises the steps of starting, two-stage compressed air cooperation, ignition and self-sustaining switching, pressure ratio matching, fault response and the like, and efficient, safe and stable operation under all working conditions can be achieved. The invention has the functions of simplifying the engine structure, improving the supercharge ratio of the gas compressor, increasing the thrust of the engine and the like. After the thrust of the engine is increased, the load capacity and the flight speed of the unmanned aerial vehicle or the cruise missile can be effectively increased.