227 results about "Jet engine" patented technology

The invention relates to a three-dimensional structure displacement stress field prediction method based on a point cloud neural operator, a medium, equipment and application. The method comprises the following steps: constructing a neural operator model for predicting a structure physical field; collecting geometric feature data of the three-dimensional structure and corresponding working condition data; training the neural operator model by using the acquired data and the query position of the internal space of the three-dimensional structure; inputting geometric feature data and working condition data corresponding to the three-dimensional structure into the trained neural operator model, and querying any spatial position in the three-dimensional structure to obtain a displacement predicted value and a stress predicted value at the corresponding position; media and equipment are realized based on the method and are applied to jet engine support displacement stress field prediction. The method can be directly used for a non-parametric complex three-dimensional structure, continuous prediction of a displacement field and a stress field of any spatial query position in the structure is achieved, the structure response prediction efficiency is remarkably improved, the calculation cost is reduced, and the method has good engineering application value.

A design method for a blade with orientation structures, a blade and a performance test method for a blade are provided, which relate to the technical field of blade design, and are universal to blades whose parameters are completely or partially known. In the design method, orientation awls are provided at a trailing edge of a blade from a variable-pitch propeller; orientation awls are provided at a trailing edge of a blade from a fixed-pitch propeller or a fan blade from a jet engine, and orientation grooves are provided on a back of the blade. Meanwhile, during providing orientation awls for either blade, the blade is stretched in a chord direction to ensure that an area of a pressure surface of a modified blade is equal to an area of a pressure surface of its original.

The high-temperature-rise combustion chamber comprises a sudden expansion diffuser, a cartridge receiver, a combined air atomizing nozzle and a flame tube and can solve the problems that in the prior art, a high-temperature-rise combustion chamber of a small and medium-sized turbojet engine is low in fuel oil supply pressure and fuel oil flow is increased, so that fuel oil atomization is poor, the combustion efficiency is low, and the combustion efficiency is low. And the service life of the flame tube is difficult to meet the requirement due to great reduction of the cooling air quantity and increase of the cooling air temperature. An overall layout form of fuel oil radial grading and axial grading is innovatively adopted, and swirlers used for assisting fuel oil atomization all adopt a beveling hole mode; space compound angles are innovatively adopted for cooling holes in the flame tube, and the overall layout form of different sizes and different hole pitches is adopted in different areas. Fuel oil atomization quality and efficient combustion under various working conditions can be improved, the service life of the flame tube is prolonged, and meanwhile the machining process of the swirler is simplified due to the fact that the swirler adopts an inclined hole mode.

The invention provides a compact medium-temperature fuel cell jet engine multi-mode split hybrid propulsion system, and aims to improve the efficiency of an aircraft propulsion system and reduce environmental pollution. Compared with a traditional engine, the engine adopts the parallel double-shaft design, flexible transmission and distribution of power are achieved through the bevel gears and the clutches, and two working modes are supported to drive a fan, namely, a shaft driving mode and an electric driving mode. And in the medium temperature mode, the engine abandons a high-pressure compressor, and air treated by a low-pressure compressor is directly sent to the cathode of the fuel cell, is heated together with mixed gas obtained after combustion of hydrogen and oxygen, and then enters the anode. The invention also considers the configuration of the galvanic pile, and adopts a hollow cylinder configuration, so as to increase the reaction area and improve the thermal management. And meanwhile, the slender structure optimizes the space utilization, so that the overall layout is more compact. The coupled lower structure allows integration of a plurality of motors, facilitating heat dissipation and maintenance. And the shell cover provides protection for the internal structure, so that the durability and the safety are enhanced. Through the innovative design, the engine has remarkable advantages in the aspects of improving the heat efficiency, reducing the oil consumption rate and reducing NOx emission, and has important significance on environmental protection and development of the aviation propulsion technology.

The invention provides a multi-mode split hybrid propulsion system of a compact high-temperature fuel cell jet engine. The problems that a traditional aviation gas turbine engine is low in heat efficiency, high in oil consumption rate and high in nitrogen oxide emission amount are solved. The device comprises a propelling system, a transmission system and a lubricating system, the propelling system adopts parallel double shafts; the two connecting shafts can transmit power through a transmission system, so that two working modes of the ducted fan are realized; and the lubricating system feeds lubricating oil to all shaft systems on the shaft. The high-temperature solid oxide fuel cell system supplies power to the motor through the direct current conversion device, then the motor drives the ducted fan, and the total efficiency and the total thrust of the hybrid propulsion system are improved. The hydrogen-rich combustion chamber and the hydrogen-poor combustion chamber are additionally arranged on the structure, so that the problem that the high-temperature solid oxide fuel cell needs a heat exchanger with a complex pipeline to preheat fuel is solved.

A propulsion assembly for an aircraft, the propulsion assembly having a jet engine having a fan casing and a central casing around a longitudinal axis and having a vertical median plane passing through the longitudinal axis, an attachment pylon having a rigid structure that takes the form of a box that has a front wall and an upper spar extending forwardly in respect to the front wall, a front engine attachment fixed between an upper area of the central casing and a lower end of the front wall, and a complementary front engine attachment fixed between an upper area of the fan casing and a front end of the upper spar.

The invention concerns an aircraft propulsion system (100) comprising a turbojet (102) having a fan casing (103) with a rear face (103a), a pylon (104) with a front part (106a) and a lower spar (106b) with a rear fitting (108) fixed to it, and a front engine attachment (150) fixed between an upper part of the fan casing (103) and the front part (106a), a front arch (152) having two arms (152b-c), each having a first end fixed to the lower spar (106b) and a second end fastened in an articulated manner to the rear face (103a), and two lateral rods (154), each having a first end fastened in an articulated manner to the rear fitting (108) and a second end fastened in an articulated manner to the rear face (103a). With such an arrangement, the forces are transmitted to the rear of the pylon.

A wearable water shoe apparatus for walking, standing and jet siding on water, includes in each of a first water shoe and a second water shoe (i) a sleeve to receive a leg of a user and (ii) a bottom section. The sleeve has a platform for standing thereon. Each pair of a plurality of pairs of jet motors is configured to eject a jet stream of water. A stabilization unit, including a tilt sensor and a plurality of gyroscopes, is configured to detect a stabilization state of at least one of the first water shoe and second water shoe. A processing unit is configured to receive output from the stabilization unit, determine a direction of movement indicated by the user and direct at least one of the pairs of motors to eject a water stream out of at least one of the water shoes based on the indicated direction.

A method for preparing an abradable panel for a rub strip of a fan case of a gas turbine engine includes forming an abradable material into a plurality of panels each comprising an abradable structure; curing the abradable material in the plurality of panels; and installing the plurality of panels into the fan case. The panels can be cured and assembled into a fully assembled rubstrip, and then the assembled rubstrip can be installed into the fan case.

An airframe is provided that can be flown under the control of an operator or an automated sensor and feedback system. The airframe can be any of a variety of known configurations with multiple rotors, fixed or retractable wings, puller or pusher propellers, or jet engines. The airframe includes at least one arm that is used to intercept and disable another airborne vehicle. In some configurations, the arm is fixed and can include fingers at its outer end. In other configurations, the arm is movable from a closed condition to an open condition. A wire or net is connected between the arm and a portion of the airframe so that the wire or net is spread open when the arm is deployed to the open condition. The wire or net are configured to disable or capture a target such as another aerial vehicle.

The invention relates to an evaporation pipe backflow combustion chamber and a jet engine, belongs to the technical field of jet engines, and solves the problem of small backflow low-speed area of a head main combustion area in a flame tube of an evaporation pipe backflow combustion chamber in the prior art. The device comprises a flame tube; a main rotational flow structure and a secondary rotational flow structure are arranged at the head of the flame tube; the main rotational flow structure is arranged at an air inlet of a head outer ring of the flame tube, and the secondary rotational flow structure is arranged at the position, close to the center line of a combustion chamber, of a head inner ring of the flame tube. The main rotational flow structure is used for generating a backflow area, and the secondary rotational flow structure plays a role in assisting in enhancing the backflow area; the main swirl vane adopts a strong swirl design, and the secondary swirl vane adopts a weak swirl design; a backflow area which is full of 1 / 3 of the total volume of the flame tube is formed in a main combustion area of the flame tube, secondary atomization mixing and staying and distribution of oil gas in the main combustion area are guaranteed, and low-working-condition combustion stability and combustion efficiency are guaranteed.

The utility model relates to a jet decontamination device which comprises a turbine jet engine and a jet pipe, the turbine jet engine is provided with a tail jet pipe, the tail jet pipe is connected with the jet pipe, the jet pipe is connected with a material guide pipe, and the material guide pipe is connected with a medicament box. High-speed airflow generated by the tail of the turbine jet engine is utilized, the jet pipe is additionally arranged at the tail, the high-speed airflow generates negative pressure through the jet pipe, a liquid or powdery decontaminating agent is sucked in, the high-speed airflow carries the decontaminating agent to decontaminate pollution in air and on the surface of an object, the spraying range is long, the spraying range is wide, and the spraying effect is good. Spraying is more uniform, a power system does not need to be additionally arranged for feeding the decontaminating agent into the turbine jet engine, and the production cost is reduced.

A thrust reverser device for an aircraft turbojet engine nacelle an electro-hydrostatic actuation system that includes a motor pump and a first actuation assembly. The first actuation assembly includes a series of at least two actuators configured to be attached to a movable cover of the thrust reverser and to move the movable cover or covers between a retracted position and a deployed position relative to the nacelle. The first electro-hydraulic actuation system further includes a volumetric flow divider with an inlet connected to the motor pump and at least two outlets, wherein each outlet is connected to an actuator. The volumetric flow divider is configured to receive a fluid and to supply the fluid to the actuators with an equal flow rate.

The present invention discloses a turbojet engine and flight equipment. The turbojet engine includes a compressor; a combustion chamber connected to the compressor; a guide, the guide including a guide body and a guide housing, the guide housing being sleeved on the outside of the guide body, and the end face of the guide body being provided with a circumferential positioning groove, the guide body being coaxially sleeved with the main shaft sleeve of the turbojet engine, and the main shaft sleeve being provided with a matching positioning groove capable of matching with the circumferential positioning groove; a circumferential positioning member, the two ends of the circumferential positioning member being respectively inserted and matched with the circumferential positioning groove and the matching positioning groove; an axial positioning member for axially positioning the guide; and a tail nozzle assembly for ejecting reverse thrust airflow. The present invention fixes the guide by means of circumferential positioning members and axial positioning members, thereby reducing the contact area between the guide and the main shaft sleeve, reducing the heat transferred to the main shaft, thereby reducing the operating temperature of the bearings on the main shaft and extending the service life of the bearings.

A cooling system for marine jet engines includes a marine jet engine assembly, a pressurized raw water inlet, a centrifugal filter, a strainer, and a hull-mounted heat exchanger. The marine jet engine assembly is configured with an engine, a drive shaft, a stator, a manifold heat exchanger, an exhaust, a water pump, and a coolant outlet so that the vessel can be functional. The pressurized raw water intake is in fluid communication with the stator. The centrifugal filter is in fluid communication with the pressurized raw water intake to supply a direct flow of raw water from the stator. The raw water is then cleaned through the strainer, discharged into a water inlet of the manifold heat exchanger, and discharged from a water outlet of the manifold heat exchanger before exiting into the exhaust. The hull-mounted heat exchanger is in fluid communication with the coolant outlet and the water pump.

The application provides a high-efficiency cooling jet engine, which comprises an inner cone and a cone tip connected with the inner cone; one end of the inner cone is provided with an inner cone front mounting edge, the other end is provided with an inner cone rear mounting edge, one end of the cone tip is provided with a cone tip mounting edge, and the cone tip mounting edge is connected with the inner cone rear mounting edge; an air film ring main body is arranged in the inner cone, one end of the air film ring main body is connected with the inner cone front mounting edge, and the other end is connected with the inner cone rear mounting edge through an air film ring baffle; a nozzle and a cooling cover are sequentially sleeved outside the inner cone, one end of the nozzle is provided with a nozzle front mounting edge, one end of the cooling cover is provided with a cooling cover front mounting edge, and the cooling cover front mounting edge is connected with the nozzle front mounting edge. The application has a double-mode cooling function, an air film ring with a large number of cooling air film holes is further designed in the cone cavity, a remarkable cooling effect can be achieved, the number of parts is small, the structure is simple, the work is reliable, the performance is excellent, the application range is wide, and the practicality is strong.

A rear part for a turbojet engine comprises a variable-geometry nozzle in which a convergent flap is provided with a lever (80) supporting a bearing roller (96) arranged axially between an upstream bearing wall (90) and a downstream bearing wall (92), which are rigidly secured to a movable part capable of moving axially on command with respect to an upstream stator structure (38) in such a way that the bearing roller is free to move with respect to the upstream and downstream bearing walls in a radial direction (R). During a downstream movement of the movable part, the upstream bearing wall pushes the lever to pivot a downstream end (46) of the convergent flap towards the longitudinal axis. Operation of the drive means for moving the movable part can thus be optimized while limiting the size and the mass of the means for controlling the flaps.

The invention relates to a turbojet afterbody comprising: – an exhaust casing (9) carrying a confluence plate (11) which surrounds a primary flow and is surrounded by a secondary flow; – an afterburner chamber (12) comprising a casing (13) and a liner (16) delimiting with this casing (13) an annular space (17) extending along an axis (AX), this liner (16) having a diameter greater than that of the confluence plate (11); – a fixed shroud (22) extending from the confluence plate (11) to the internal liner (16), this fixed shroud (22) comprising windows (27); – a movable ferrule (23) running along the fixed ferrule (22) and which comprises windows (28), this ferrule (23) being movable in rotation around the axis (AX) between an open position in which the windows (27) of the fixed ferrule are open and a closed position in which the windows (27) of the fixed ferrule (22) are closed by the movable ferrule (23).Figure for short: Figure 2.

FIELD: rocket science.SUBSTANCE: invention concerns the design of a ramjet engine. The ramjet engine body is a structure consisting of several layers, including: an internal heat-protection layer made of a composite material based on a thread frame of heat-resistant fibers impregnated with an elastic polymer binder, a power body based on a high-temperature polymer composite material and an external heat-protection layer made of a composite material based on an elastic polymer binder and a finely dispersed mineral filler. Thanks to the use of layers of external and internal thermal protection, it becomes possible to use a polymer composite material for the power casing.EFFECT: reducing the labor intensity of manufacturing and weight of the product and increased energy-mass perfection of the propulsion system.1 cl, 2 dwg

The invention discloses coal mine underground space operation rapid escape equipment and method.The coal mine underground space operation rapid escape equipment comprises an escape equipment main body pipeline, two escape rings are arranged at certain intervals on the escape equipment main body pipeline, a supporting column is arranged between the two escape rings, an oil conveying pipeline is arranged in the supporting column, and one end of the oil conveying pipeline is connected with an upper escape ring; one end of the upper escape ring is connected with the escape equipment main body pipeline, the other end of the upper escape ring is connected with the micro jet engine, the micro jet engine is arranged at the bottom end of the lower escape ring, LED lamps are arranged at the top end of the upper escape ring and the bottom end of the lower escape ring, rolling pulleys are arranged between the escape ring and the escape equipment main body pipeline, and a telescopic system is arranged in the escape ring. According to the escape equipment, rapid escape of underground operation can be achieved, the survival probability of operating personnel is increased, meanwhile, the equipment can be recycled, the economical efficiency and feasibility of the underground operation escape equipment are greatly improved, the weak feasibility of underground space operation safety equipment is improved, and the safety of the underground space operation safety equipment is improved. And the shortage of an escape protection device for underground space operation is made up.

A propulsion assembly for an aircraft, the propulsion assembly having a jet engine having a fan casing and a central casing around a longitudinal axis and having a vertical median plane passing through the longitudinal axis, an attachment pylon having a rigid structure that takes the form of a box that has a front wall and an upper spar extending forwardly in respect to the front wall, a front engine attachment fixed between an upper area of the central casing and a lower end of the front wall, and a complementary front engine attachment fixed between an upper area of the fan casing and a front end of the upper spar.

An aircraft, an aircraft bleed air system, and a method of providing thrust for an aircraft. The bleed air system includes a bleed air turbine configured to receive at least a first portion of bleed air from a jet engine and to deliver the at least first portion of bleed air to a cabin, a cabin bleed air compressor operably coupled to the bleed air turbine and configured to receive cabin bleed air from the cabin, and a cabin bleed air nozzle configured to receive cabin bleed air from the cabin bleed air compressor. The aircraft includes a fuselage, a wing, a jet engine, and the bleed air system. The method includes the steps of expanding at least a first portion of bleed air from a jet engine to drive a cabin bleed air compressor, compressing cabin bleed air from a cabin with the cabin bleed air compressor, and ejecting the cabin bleed air into an environment.

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).

The present invention relates to an exhaust nozzle for fluid thrust vectoring comprising an elongated channel for directing exhaust gases having first and second ends. The elongated channel has an inlet at the first end with a continuously curved cross-sectional shape and an outlet at the second end with a polygonal cross-sectional shape. In a preferred embodiment, the exhaust nozzle has an outlet of rectangular cross-section. A jet engine and an aircraft having the exhaust nozzle are also provided.

The invention discloses a water ramjet injector, which is characterized in that a core component is a centrifugal two-stage atomizing nozzle, and a two-stage duct design is adopted in the centrifugal two-stage atomizing nozzle: the first stage is a communicating cavity, high-pressure water flow enters the first atomizing cavity and then is accelerated by the communicating cavity, a strong shearing effect is generated due to sudden change of the aperture, and high-Reynolds-number turbulent flow, namely first-stage atomization, is formed; and then the fluid enters the second atomization cavity and interacts with the gas medium at an outlet, and the liquid film is further crushed into micron-sized fog drops, namely second-stage atomization. The nozzles are restrained to the inner wall of the combustion chamber through threads in a circumferential axial symmetry mode, and the jet flow axes and angles of the multiple nozzles are reasonably arranged, so that atomized water media achieve self-impact collision in the center area of the combustion chamber. In the collision process, through momentum accumulation and turbulence disturbance, liquid drops are promoted to be subjected to secondary crushing, the atomization uniformity and the mixing efficiency are remarkably improved, and the method is suitable for large-scale industrial use and popularization.

This disclosure relates to a monitoring system and method for monitoring a jet engine in transit. The system includes a processor, transceiver, sensor and digital storage media for measuring environmental variables with respect to thresholds. A sensor probe is also disclosed for monitoring the status of a jet engine enclosed in sheeting.

A matrix for a heat exchanger to exchange heat between a first fluid and a second fluid, the first fluid being for instance air and the second fluid being for instance oil. The matrix comprises: a channel for the first fluid. an array of passages for the second fluid, the passages extending in the channel. The array supports at least two cooling fins. The matrix is made by a process of additive manufacturing. The fins are inclined with respect to each other along the direction of the flow of the first fluid. The array defines rectangular corridors for the first fluid.

Combined jet engine with a compressor driven by an internal combustion engine, comprising an internal combustion engine, a control gearbox, a coupling device, an air compressor, a mixing chamber with a central body, an outlet device - jet nozzle and an automatic control system, characterized in that the air compressor (2.7 - 2.8) is driven by an internal combustion engine (1), wherein the internal combustion engine is arranged upstream of the air compressor in the direction of flow, the air compressor is single-stage and has a front power body with inlet guide vanes (2.6), wherein the rotor main bearing of the air compressor (2.7 - 2.8) with a support bearing is housed inside the body, a mixing chamber (2.11) with a central body (2.10) in the form of a profiled, perforated jacket is arranged behind the air compressor, and an uncontrolled jet nozzle is arranged behind the mixing chamber;

The invention discloses a jet-propelled multi-energy efficient aircraft which comprises an aircraft body, main wings are symmetrically arranged on the two sides of the aircraft body, at least one jet-propelled engine is arranged on the main wings, and the jet-propelled engine is provided with a nozzle; the auxiliary wings are symmetrically arranged on the two sides of the aircraft body and located on the rear sides of the main wings. The airflow conversion cabins are arranged on the auxiliary wings and used for capturing high-speed airflow sprayed out of the nozzles, accelerating the high-speed airflow and then spraying the high-speed airflow out of the output ends of the airflow conversion cabins so as to generate additional reverse thrust for pushing the aircraft main body to move forwards; main thrust is provided for the aircraft body through the jet engine, aerodynamic loads are balanced through the symmetrically-designed auxiliary wings, originally-wasted jet airflow kinetic energy is captured through the airflow conversion cabin, additional reverse thrust is generated by accelerating airflow, advancing power of the aircraft is supplemented, and secondary utilization of the jet airflow kinetic energy is achieved. Thrust needing to be output by a main engine is reduced, fuel consumption per unit mileage is reduced, and emission of carbon dioxide is reduced synchronously.

The application discloses a rotor parameter decoupling estimation method for an unmanned aerial vehicle based on an OMP algorithm, and the method comprises the following steps: obtaining a rotor unmanned aerial vehicle radar echo, and completing signal sampling, pulse compression and other pretreatments; constructing an echo matrix and performing motion compensation and filtering on a rigid body part; estimating a target rotor rotating speed according to the jet engine modulation characteristics of the rotor echo; constructing a phase matching item and performing initial phase estimation on a target rotor blade; designing an overcomplete dictionary based on the phase matching item, and utilizing the OMP algorithm to reconstruct the sparse distribution of a target rotor scattering point. The application can realize accurate estimation on the rotating speed, the number and the length of the target rotor blade under the condition of echo azimuth under-sampling, realizes parameter decoupling, effectively reduces the calculation complexity, reduces the algorithm running time, and makes up for the shortcomings of the prior art.