183 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 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.

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.

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.

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

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 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.

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.

The invention relates to the technical field of welding, and discloses a large-size steel structure welding preheating device which comprises a turbofan jet, one end of the turbofan jet is an air inlet channel, the other end of the turbofan jet is a rear combustion chamber, the turbofan jet is sleeved with a vortex sleeve, the front end of the air inlet channel is provided with a nitrogen compressor, and one end of the vortex sleeve is provided with a cold air discharge pipe. The cold air discharging pipe and the air inlet channel are located on the same side, and the hot air discharging pipe and the rear combustion chamber are located on the same side. Compared with the prior art, the device has the advantages that the turbofan jet generates high-temperature fuel gas flow, large-flow and high-flow-speed hot gas flow injection is achieved, the preheating range is wider, cold nitrogen is mixed into the gas inlet channel through the nitrogen compressor, the oxygen content of a combustion area is reduced, and the oxidation risk of a preheating area is reduced; the hot gas discharge pipe forms a low-oxygen hot nitrogen layer on the periphery of the high-temperature tail gas, so that oxidation is further inhibited while heat loss is reduced.

The invention relates to a transport tool (40) configured to be temporarily mounted on a thrust reverser of an aircraft turbojet engine nacelle of a propulsion assembly, the tool (40) being intended to connect a front frame and a rear frame secured to one another. The transport tool (40) comprises: - at least one main portion (41, 42) extending along a first axis (X); and - at least one signalling blade (44) connected to the main portion (42) and extending in a use position along a second axis (Y) inclined with respect to the first axis (X).

A rear section for a turbojet engine includes a variable geometry nozzle in which a converging flap is provided with a lever (80) supporting a support roller (96) axially arranged between an upstream support wall (90) and a downstream support wall (92), the upstream and downstream support walls being rigidly fixed to a movable section capable of axial movement relative to an upstream stator structure (38) upon command, such that the support roller moves freely in the radial direction (R) relative to the upstream and downstream support walls. During downstream movement of the movable section, the upstream support wall pushes the lever to pivot the downstream end (46) of the converging flap toward the longitudinal axis. Therefore, the operation of the drive mechanism for moving the movable section can be optimized while limiting the size and mass of the device used to control the flap.

The invention discloses a valveless pulsating jet engine with an oval air inlet channel. The valveless pulsating jet engine is composed of a combustion chamber, the air inlet channel, a tail pipe, an oil nozzle and the like. An engine air inlet channel is of an oval flow channel structure, the air inlet channel and a tail pipe are arranged on the rear wall face of a combustion chamber, an oil nozzle is arranged on the rear half portion of the combustion chamber, and a nozzle directly faces an air inlet. The air inlet channel has the advantages that under the constraint of the limited engine section size, a circular flow channel is changed into an oval flow channel by changing the structural form of the air inlet channel, the air inlet amount of an engine is effectively increased, and therefore the thrust of the engine is improved.

The invention relates to a method for developing a scanning process, a method for acquiring three-dimensional data of a blade and a system for acquiring position data of different types of blades. Techniques are provided for acquiring data from an aircraft component, such as from a turbine blade for use in a jet engine. A specially designed fixture with a reference point is configured to hold a blade to be scanned. A 3D scanner is used to acquire position data of the blade at high resolution, and the data is used to generate a 3D model. The 3D model may be used to verify that a component of, for example, an aircraft has been manufactured as designed.

The application provides a high-compact heat management type jet engine multi-mode split hybrid propulsion system. The propulsion system has a parallel double-shaft structure, shaft work of a horizontal shaft I is transmitted to a vertical shaft through a bevel gear set I, and the vertical shaft transmits the transmitted shaft work to a horizontal shaft II through a bevel gear set II. The horizontal shaft II rotates to drive a fan to work and generate propulsion work. Three groups of low-temperature proton exchange membrane fuel cell stacks are coaxially and parallelly arranged on the horizontal shaft I, and four electric motors are coaxially and serially arranged on the horizontal shaft II. The propulsion system has two different working modes. One is hybrid power propulsion, in which the clutch does not work, and the fan can be driven by the transmitted shaft work or by the electric motor. The other is pure electric propulsion, in which the clutch works, the vertical shaft is disconnected, and the horizontal shaft I outputs useless work. The fan can only be driven by the horizontal shaft II driven by the electric motor.

The invention relates to the weapons industry, in particular to firearms. The firearm comprises a barrel (1) with a muzzle cut (2) and is provided with an external jet engine (9) configured to generate a jet of gases (10) coaxial with the barrel axis beyond the muzzle cut (2) in the direction of firing. The invention also concerns a projectile (6) for a firearm, comprising a cavity (18) containing a flowable substance (19) and one or more channels (20) communicating with the cavity (18) and having outlet openings (22) on an outer surface of the projectile (6). The cavity (18) is provided with a piston (23) capable of rectilinear movement and adapted to receive the impact of gases of the propellant charge (5). According to the invention, the projectile (6) has two alternative configurations. In the first configuration, a surface of the cavity (18) has one or a plurality of spiral -shaped protrusions (24) made along the axis of rotation of the projectile (6) and adapted to interact with the piston (23). In the second configuration, the piston (23) is made as a stepped piston consisting of several cylindrical surfaces of different diameters corresponding to the cylindrical surfaces of the cavity (18).