21 results about "Jet engine" patented technology

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

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.

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 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.

A Ramjet solid fuel comprising up to 50% metal by weight selected from the group consisting of aluminum, boron, boron carbide, magnesium, titanium, and titanium boride; between 2.0 and 9.0 weight percent fluorinated combustion aids selected from the group consisting of 3-hydroxybenzo trifluoride (3HBTF), hydroxyfluoro-2-propanol (HF2P), polytetrafluoroethylene, and trifluoroethanol (TFE); and up to 48% by weight of the remaining component being a hydroxyl terminated, epoxidized polybutadiene polymer (Poly BD 605E) (HO—CH2-CH═CH—CH2-CH2-CH—O—CH—CH2-CH2-CH—(CH═CH2)-CH2-CH═CH—CH2-CH2-CH—O—CH—CH2-OH).

Improved cryogenic fluid supply device for an aeronautical turbojet engine. Cryogenic fluid supply device (10) for an aeronautical turbojet engine (1), a tank (20) storing the fluid in liquid form, the device (10) comprising a first heat exchanger (30) capable of first heating the fluid drawn from the tank (20) via a sampling line (11) connecting the tank (20) to the first heat exchanger (30), and a main line (12) connected in a loop to the first heat exchanger (30), the device (10) comprising a bypass channel (14) comprising a first branch (141) branching from the main line (12) at a bypass point (140), to a mixer (50) capable of mixing a liquid phase and a gaseous phase of the fluid, and a second branch (142) connecting the mixer (50) to the sampling line (11), a bypass valve (60) disposed downstream of the bypass point. (140),and movable between a first position allowing a deviation of the fluid flow via the bypass channel (14), and a second position limiting this deviation. Fig. 3.

This invention discloses a method for measuring the temperature distribution of aero-engine exhaust plumes. Specifically, a rotatable robotic arm with guide rails is used. A color industrial CCD camera is mounted at the front end of the robotic arm. This robotic arm ensures that the CCD camera can capture images of the aero-engine exhaust plume 360° equidistantly around its central axis, obtaining radiation information from various angles. A circular temperature model is established for the cross-section of the exhaust plume. Radiation information is extracted from the multi-angle images of the exhaust plume. A large set of equations is constructed using the model and solved using the LSQR method to obtain the three-dimensional temperature distribution of the exhaust plume. This invention provides a method for measuring the temperature distribution of aero-engine exhaust plumes. This method can collect information from as many angles as possible (with a diameter reaching the meter level) of the aero-engine exhaust plume to ensure the accuracy of the inversion calculation. Experimental verification shows that the method of this invention can effectively measure the temperature field distribution of aero-engine exhaust plumes.

The invention relates to a turbojet engine including a shaft surrounded by a low-pressure rotor surrounded by a coaxial and independent high-pressure spool, this turbojet engine including from upstream to downstream: a fan driven by the shaft; a low-pressure compressor carried by the rotor; an inter-compressor casing; - a high-pressure compressor and a high-pressure turbine carried by the high-pressure spool; an inter-turbine casing; a low-pressure turbine carried by the rotor; an exhaust casing; this turbojet engine including: an upstream rotor bearing carried by the inter-compressor casing; a downstream rotor bearing carried by the exhaust casing; a gearbox downstream of the downstream bearing and through which the rotor drives the shaft; a downstream shaft bearing downstream of the downstream rotor bearing.

The invention relates to a high-speed exhaust nozzle, in particular of a fan or a jet engine, with a characteristic cross-sectional shape that forms a closed contour formed by six continuously connected segments, these segments comprising alternating rectilinear and arcuate segments, shaped symmetrically relative to the axis passing through the centre of the cross-section. Preferably, the cross-sectional shape is defined by a mathematical function that determines the form of an area of even distribution of pressures exerted by a stream of gases flowing through the nozzle according to the invention. The nozzle is characterised by the fact that its cross-section has a shape defined by the following relationships: (I) where the functions γk, k=1,2,3,4,5,6 have the following form (II) for t∈[0;1]. Preferably, at the midpoint of the segments defined by the function (I) for γ2; γ4; γ6, there are gas jet guide vanes pointing towards the centre of the nozzle.

An aircraft propulsion assembly having a jet engine with a front casing, a primary structure with a lower spar, two outer rods, wherein each one has a rear end mounted in an articulated manner on the lower spar and a front end mounted in an articulated manner on the front casing, and two inner rods, wherein each one has a rear end mounted in an articulated manner on the lower spar and a front end mounted in an articulated manner on the front casing. Also an aircraft with such an assembly.

Wingtip shields are described. In one embodiment, an aircraft structure includes an airfoil having a high-pressure side and a low-pressure side, where the airfoil is configured for attaching to a main body. The wingtip shield includes an inner surface facing a high-pressure side of an airfoil; an outer surface configured opposite from the inner surface. An air mover is configured at least partially within the wingtip shield. In different embodiments, the air mover is a fan, a mechanically driven propeller, or a jet engine. In some embodiments, the aircraft structure includes a shaft coupled with the air mover, a bearing configured to support the shaft, and a motor configured to rotate the shaft.

The invention relates to a rear assembly (1) of a turbojet nacelle comprising: a thrust reverser including a rear fixed structure; a combustion gas ejection nozzle (3) extending from the rear fixed structure; and a suspension pylon (4) to which the rear fixed structure is attached. The rear assembly includes a fire-sealing device comprising a fire-resistant strip (6) made of a flexible, fire-resistant material, and having opposing first longitudinal edges (61) and second longitudinal edges (62). The first longitudinal edge of the fire-sealing device is fixed to at least one of the suspension pylon and the rear fixed structure, and the second longitudinal edge is fixed to the combustion gas ejection nozzle, with the fire-resistant strip extending at least below the suspension pylon. Abstract figure: Figure 3

A dual flow diffuser includes a core flow diffuser and a bypass flow diffuser, where the core flow diffuser and bypass flow diffuser are three-dimensional (3D) printed as a unitary unit that can also contain compressor components. In some examples, the core flow diffuser and bypass flow diffuser can include hollow blades to allow engine service components and / or instrumentation to pass through the diffuser. The core flow diffuser and bypass flow diffuser can include radial and / or circumferential passages to allow fluid, such as lubricant, to flow therethrough. A surge chamber can be provided along with an adjacent through hole. An inter-stage bypass bleed device can be included having at least two circumferential chambers configured to achieve a substantially uniform pressure distribution at the bleed port. In some examples, the unitary unit can be additively 3D printed with metal.

The invention relates to jet engines, and more particularly to pulse detonation engines. The desired technical result, namely increased energy efficiency and reliability, is achieved in an engine having a combustion chamber configured in the form of a detonation resonator that opens into an exhaust nozzle. The resonator is an aspherical reflector that is symmetrical about the axis of the engine. A gaseous fuel and a gaseous oxidizer are used in the engine, together with a single-step combustion process. If a hydrocarbon fuel is to be used, the fuel duct is provided with a pyrolysis chamber. Pyrolysis is achieved as a result of the fuel flow coming into contact with the heated rear side of the reflector. A mixture of fuel and oxidizer is fed into the combustion chamber through an annular supersonic injection system. To initiate detonation, the engine can be provided with a detonation initiator in the form of a tube which is blind at the distal end and open at the end disposed in the combustion chamber, said tube lying along the axis of the engine. The detonation products flowing through the exhaust nozzle generate a thrust which propels the engine in the opposite direction. This pulse detonation engine design is simple and reliable and provides for the highly efficient generation of propulsive thrust force by maintaining cyclic resonant detonation at a frequency of around 10 kHz.

This invention relates to an unmanned surface vessel (USV) with a rapid maneuvering and anti-interference steering assistance system. By installing a jet engine and a simple engine turntable on the mast, aerodynamic thrust is provided to the hull in all directions, transferring steering power from underwater to above water, thus overcoming the influence of water flow resistance. The turntable can adjust to the target angle within milliseconds, rapidly changing the direction of steering power; the jet engine's power adjustment time is also very short, quickly changing the direction and magnitude of the steering power on the hull side. Based on a novel USV platform with added equipment, this invention has developed a control allocation strategy to maximize the efficiency of the two power systems. Compared with traditional USV systems, this invention changes the power source for heading control from underwater to above water, and attitude changes from servo deflection to jet power, eliminating the lag caused by water flow resistance and providing a rapid response capability to change heading. This solves many problems related to USV heading control and crosswind resistance.

This application discloses a microwave-based jet engine afterburner, belonging to the field of thermomechanical technology. It includes a microwave heating mechanism, whose microwave generator comprises a cathode electron source, a heating component, a ring-shaped metal anode, an electromagnet, and a DC power supply component. The cathode electron source is located at the center of the ring-shaped metal anode. The heating component causes the cathode electron source to emit free electrons. The DC power supply component supplies power to the cathode electron source and the ring-shaped metal anode to form a DC electric field and supplies power to the electromagnet to generate a magnetic field orthogonal to the DC electric field. The strip-shaped groove within the ring-shaped metal anode serves as a microwave resonant cavity. The movement of free electrons excites microwaves, which are used to heat the exhaust gas from the engine's main combustion chamber. The microwave generator is coaxially positioned within the gas flow channel. This device replaces traditional afterburning with microwave heating, offering faster heating speed, a more compact structure, and improved thermal efficiency of the afterburner, solving the problems of easy flameout and low combustion efficiency in traditional afterburner combustion chambers.

The utility model discloses a kind of unmanned aerial vehicle for realizing vertical take-off and attitude control using small turbojet engine, small turbojet engine is installed at wing tip of wing, small turbojet engine is installed at the back of fuselage (or in front of fuselage) two sides;Small turbojet engine is made to produce upward thrust by vertical rotation 90 ° using turbine worm drive mechanism or motor with self-locking holding torque, and unmanned aerial vehicle vertical take-off is realized;Unmanned aerial vehicle has aileron installed at the trailing edge of left and right wing tips, and the differential deflection of aileron makes unmanned aerial vehicle tilt rotation around longitudinal axis, and the direction of thrust of small turbojet engine is deflected using turbine worm drive mechanism, to realize the flight state stabilization and flight attitude control of unmanned aerial vehicle, so that unmanned aerial vehicle has good maneuverability;Horizontal stabilizer and elevator are omitted, vertical stabilizer and rudder are omitted, so as to reduce flight resistance, which is conducive to improving the flight speed and effective load of unmanned aerial vehicle.

The application discloses a kind of combustion optimization control system and method for dual-fuel micro-injection engine, it is related to dual-fuel engine technical field, including working condition monitoring module, pilot oil compensation control module and fuel execution module, the pilot oil compensation control module is used for based on real-time working condition dynamic correction pilot oil injection parameter, the pilot oil compensation control module includes: injection parameter prediction unit, compensation amount calculation unit and injection correction unit.The application is installed with pilot oil compensation control module, adopts support vector machine model to predict reference parameter and dynamically compensate adaptive controller combination, realize millisecond level pilot oil injection correction, in fuel-gas mode switching process, combustion phase offset amount is controlled within ±0.8°CA, and speed fluctuation rate is reduced to below 5%, effectively eliminates torque mutation and the risk of flameout.

The application discloses a valveless pulse jet engine quick starting device and method, which comprises a valveless pulse jet engine, an igniter, a blowing system and an oil supply system; after alcohol is filled in the alcohol tank and the engine starts, the igniter starts to work first, alcohol is pressed into the combustion chamber by fuel, the blowing system blows high-pressure air into the combustion chamber from the air inlet, and the blowing system and the igniter are closed in turn after a period of time, so that the engine can be started reliably. The method of the application can improve the starting reliability of the engine and greatly shorten the starting time of the engine by using alcohol combustion assisting preheating and high-pressure air blowing excitation.

This invention provides a direct-injection swirl-type mixing nozzle for a water-jet engine and a water-jet engine itself, relating to the field of water-jet engine technology, to solve one of the technical problems of low nozzle flow rate, poor atomization effect, and short penetration depth in water-jet engines. The direct-injection swirl-type mixing nozzle for a water-jet engine includes a core and a cavity enclosed by a shell, the shell having opposing inlet and outlet ends; along the direction from the inlet end to the outlet end, the cavity sequentially has a first cavity section and a second cavity section, the core being disposed in the first cavity section, and the core having at least one channel for water flow. This nozzle, used in water-jet engines, can improve atomization effect while maintaining a large flow rate.

This utility model proposes a jet engine base for aluminum surface processing, including a support column and a paint pump. One end of the support column has a tray, and the other end has a base. One side of the base has a first wheel and a second wheel, and the other side has a sliding groove. Inside the sliding groove are a first block and a second block. The first block has an annular groove around its perimeter, and the second block has a through groove inside. A limiting block is located on one side of the annular groove, and a lifting block is located in the through groove. The bottom of the limiting block has a recessed keyway, and one side of the lifting block has a protruding key. The bottom surface of the lifting block has multiple elastic blocks, and the top surface has a control rod. The key of the lifting block can pass through the keyway. The paint pump can be inserted inside the sliding groove. The bottom of the paint pump has a protruding push rod that can contact one end of the limiting block. The limiting block can form a positioning and annular movable structure. The beneficial effects of this utility model are: its structural design is simple and reasonable, further improving the assembly structure of the base and paint pump, and facilitating disassembly and installation.