140 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.

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 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.

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.

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).

The invention relates to a thrust reverser ring of an aircraft jet engine extending along a longitudinal axis. The ring comprises thrust reverser modules each extending in a longitudinal direction (D) parallel to a longitudinal axis, the thrust reverser modules being circumferentially distributed around the axis to form the ring. A thrust reverser module (10) is arranged between and mechanically independent of two adjacent thrust reverser modules, said thrust reverser module (10) comprising: a spar (12) extending in a longitudinal direction (D) and comprising two opposite ends (12c, 12d) spaced apart from each other in the longitudinal direction, each end being configured to attach the spar to an aircraft nacelle; and a plurality of transverse blade elements (14a, 14b) distributed along the spar, extending from the spar.

The invention relates to a device (30) which can be used for cleaning a jet engine (1) and to an arrangement (10) and method for cleaning a jet engine. The device (30) is used to reduce the proportion of foreign liquid in the engine oil that enters the engine oil system (2) of the jet engine (1) during the cleaning process of the jet engine (1) and comprises supply and discharge lines (31, 32) for the transparent serial intermediate connection of the device (30) into the oil circuit of the engine oil system (2), and a treatment unit (33), which can thus be temporarily integrated into the oil circuit of the engine oil system (2) in order to reduce the foreign liquid content in the engine oil flowing through the treatment unit (33). The arrangement (10) comprises a unit (20) for introducing a liquid cleaning medium into the jet engine (1) and a device (30) according to the invention. The method relates to the operation of the arrangement (10) for cleaning a jet engine (1).

The application provides a compact high-temperature fuel cell jet engine multi-mode split hybrid propulsion system, which is beneficial to solve the problems of low thermal efficiency, high fuel consumption and high nitrogen oxide emission of traditional aviation gas turbine engines. The system comprises a propulsion system, a transmission system and a lubrication system; the propulsion system adopts parallel double shafts; the two connecting shafts can transmit power through the transmission system to realize two working modes of the ducted fan; the lubrication system sends lubricating oil to all shafts on the shaft. The high-temperature solid oxide fuel cell system is used to supply power to the motor through the direct current conversion device, and then the motor drives the ducted fan, so that the total efficiency and total thrust of the hybrid propulsion system are improved. The hydrogen-rich combustion chamber and the hydrogen-lean combustion chamber are added to the structure, so that the problem of fuel preheating of the high-temperature solid oxide fuel cell through the heat exchanger with complex pipelines is solved.

Output signals from ground power units can be combined to supply more energy for aircraft to use for starting a jet engine. The combining can be possible when the outputs from the ground power units are synchronized. In a synchronization process, a first ground power unit may be started so that its first output signal serves as a reference for the synchronization. Then, an interconnect on a second ground power unit can be toggled to start and stop the generation of a second output signal at different times until the signals are in phase (i.e., synchronized).

Aircraft turbojet engine comprising at least one structural panel connected to a structure by at least one removable connection and at least one articulation The invention relates to an aircraft turbojet engine comprising: a structure (52), at least one structural panel (60), at least one removable connection configured to occupy a mounted state in which it immobilizes the structural panel (60) relative to the structure as well as a disassembled state in which the structural panel (60) can be moved away from the structure (52), at least one articulation (70) connecting the structural panel (60) and the structure (52) and allowing, when the removable connection is in the disassembled state, the structural panel (60) to move between positions close to and apart from the structure (52). This articulation (70) makes it possible to support the structural panel (60) and to move it away from the structure (52) when the removable connection is disassembled. Figure 7

The utility model provides a turbojet fire extinguisher with a telescopic connecting rod, which is characterized in that an oil path quick joint (7) at the rear part of a handle shell (3) is communicated with an oil path outlet (10) at the front part of the handle shell (3) sequentially through a turbojet oil pump (8) and an oil filter (9) in the handle shell (3), and an air inlet (17) of a turbojet engine (16) is opposite to an air inlet hole (12); a telescopic connecting rod is installed between an oil way outlet (10) in the front of a handle shell (3) and the rear end of a turbojet air duct (11), a branch pipe body I and a branch pipe body II of the telescopic connecting rod are an oil outlet pipe I (19) and an oil outlet pipe II (20) respectively, and the oil way outlet (10) is communicated with an oil supply port (21) of a turbojet engine (16) through the oil outlet pipe I (19) and the oil outlet pipe II (20). And a locking device is arranged between the oil outlet pipe I (19) and the oil outlet pipe II (20). According to the utility model, the energy waste is reduced in the fire extinguishing process, and the use safety is improved.

The application discloses a solid hydrogen storage device, which comprises a cylinder, a hydrogen outlet arranged at the top of the cylinder, a tail gas inlet arranged at the bottom of the cylinder, an inner container arranged in the cylinder, a solid hydrogen storage material arranged in the inner container, a metal pipe arranged at the bottom of the inner container, one end of the metal pipe being communicated with the tail gas inlet, an electric heating wire arranged outside the metal pipe, and a heat insulation layer arranged between the cylinder and the inner container. The application further provides an ammonia-hydrogen composite jet engine based on the solid hydrogen storage device and a control method.

Disclosed is an aircraft having a rocket-based combined cycle (RBCC) propulsion system including a rotating detonation rocket engine (RDRE) and an airbreathing jet engine. The RDRE is located within a throat area of the jet engine and is configured to be cooled by air bypassed from upstream of the jet engine.