304results about "Jet type power plants" patented technology

A system for attaching a front engine of an aircraft, having: a vertical median plane, an engine pylon with a nose having a main female clevis, and a cylindrical housing around a vertical axis, a main link fastened to a front casing of the engine by a secondary ball-joint connection about a secondary axis perpendicular to the median plane by a secondary shaft, a main shaft, perpendicular to the median plane and forming a main ball-joint connection of the main link in the main female clevis about a main axis, the main and secondary axes in the plane perpendicular to the median plane, and an outer cylinder, coaxial with the vertical axis and a first end of which fastened in the cylindrical housing and a second end of which mounted, via an annular linear connection, around the vertical axis with respect to the front casing.

The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.

An aircraft propulsion assembly including an engine assembly that includes front and rear parts offset along a longitudinal axis, a primary structure of a pylon, at least one supporting structure that extends along the longitudinal axis, separate from the rear part and connected to the front part of the engine assembly, an engine-attachment system that includes at least one front attachment directly connecting the engine assembly and the primary structure, and at least one rear attachment offset toward the rear relative to the front attachment connecting the primary structure and the supporting structure.

An aircraft may comprise propulsors. The propulsors may be integrated into an array of propulsors. An array of propulsors may be integrated in an airfoil of the aircraft. The aircraft may include a flow turning control system. The flow turning control system may include an end wall and one or more movable flaps. The flow turning control system may enable selective flow turning of the flow of propelled by the array of propulsors. The selective flow turning may include actuating the one or more movable flaps. The selective flow turning may achieve flow turning of over a trailing edge of a movable flap.

Truss-braced wing aircraft engine mounting and associated systems are disclosed. An example aircraft includes a fuselage including a central structural section, a wing extending from the fuselage, a truss extending from the central structural section and coupled to the wing, a pylon coupled to the central structural section, and an engine coupled to the pylon, wherein the pylon is positioned at an angle substantially 45 degrees from a horizontal plane and a vertical plane.

A propulsion system for an aircraft can include an electric power source and an propulsion assembly having a propulsor. An electric power bus can electrically connect the electric power source to the propulsion assembly. The electric power source can be configured to provide electrical power to the electric power bus. An inverter converter controller can be positioned along the electric power bus and can be electrically connected to the electric power source at a location downstream of the electric power source and upstream of the electric propulsion assembly.

A propulsion assembly having an engine and an engine attachment which connects the engine and a primary structure of a pylon and which comprises a removable connection connecting a transverse beam and the primary structure. This removable connection has a first contact surface integral with the transverse beam, a second contact surface integral with the primary structure, at least one connection element keeping the first and second contact surfaces pressed against one another, at least one system for transmission of transverse forces between the engine and the primary structure and at least one waiting fail-safe connection distinct from each connection element and configured to generate an additional path for forces only in the event of malfunction of an element of the first removable connection.

An assembly for an aircraft having an upper fitting fastened to a structure of the wing of the aircraft, a lower fitting fastened to a pressure-side panel of the wing, an engine pylon with two lateral panels mounted so as to articulate on the lower fitting, a starboard-side link rod and a port-side link rod mounted so as to articulate between the lateral panel that is on the same side and the upper fitting, and a stud having a proximal end secured to a rear rib of the engine pylon and a distal end mounted in an aperture of the lower fitting through an annular linear connection.

An aircraft may comprise propulsors. The propulsors may be integrated into an array of propulsors. An array of propulsors may be integrated in an airfoil of the aircraft. The aircraft may include a flow turning control system. The flow turning control system may include an end wall and one or more movable flaps. The flow turning control system may enable selective flow turning of the flow of propelled by the array of propulsors. The selective flow turning may include actuating the one or more movable flaps. The selective flow turning may achieve flow turning of over a trailing edge of a movable flap.

The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.

An assembly comprising a longitudinal member of an aircraft wing, where the longitudinal member has a contact wall which is penetrated by first bores, each passing through the contact wall parallel to a longitudinal direction, an engine pylon which is arranged at the end of the longitudinal member in the longitudinal direction and having an upper longitudinal member which, for each first bore, is penetrated by a second bore which is aligned with the first bore, and a nut and a screw for each pair of a first bore and a second bore, which are aligned, where the screw is screwed into the nut successively via the first bore and the second bore, sandwiching the upper longitudinal member and the contact wall.

The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.

The present invention provides an infrared enhancer with controllable radiant power, which relates to the field of unmanned aerial vehicles, and solves the problem that it is difficult to control the radiant power of the infrared enhancer; the device comprises a support component, a turbojet engine, which is fixed inside the support component; an installing adapter, which is used to connect the product to be installed, which is provided at one end of the support component, and is located on the side of the flame erupted from an exhaust nozzle of the turbojet engine; a heating cone, which is fixed to one end of the installing adapter and toward the exhaust nozzle of the turbojet engine, wherein infrared enhanced coating is coated outside the heating cone, and a heat insulation device is provided inside the heating cone; the present invention uses the flame at the exhaust nozzle of the turbojet engine to heat the heating cone, the flame temperature at the exhaust nozzle can be controlled by controlling the size of the operating state of the engine, and the heating cone is heated to form a stable infrared radiation source to enhance the infrared characteristics of the target; the infrared point source characteristics of an aviation jet engine can be simulated, and the present invention is suitable for a larger flight speed domain and an altitude domain.

A hydrogen fuelled aircraft comprising a fuselage and a wing, and at least one fuel tank. The hydrogen fuel tank is mounted to the wing, wherein the fuel tank is mounted above the wing and spaced from an upper surface of the wing by a tank mounting arrangement which does not interfere with the wing upper surface. [Use Figure 2]

An aircraft power generation system includes: a turbo fan engine including a variable stator vane which is located upstream of a fan and whose installation angle is variable and a vane actuator that drives the variable stator vane; a power generator mechanically connected to a low-pressure shaft of the engine; and processing circuitry configured to control the vane actuator. The processing circuitry is configured to, when it is determined that the power generator generates electric power during an operation of the turbo fan engine, execute an electric power generating mode of controlling the vane actuator to change the installation angle of the variable stator vane such that the flow rate of air flowing through the fan decreases.

An apparatus for a sensor for monitoring the health of an electric power train (12) is disclosed herein. A camera (100) includes a camera housing (202) having an imaging end (208) with an end wall (210). A lens (220) positions at or within the end wall (210) and an air circuit (230) passing through the end wall (210) and positioned to direct a supply of air (A) along or at the lens (220). The camera (100) images the electric power train (12) or electric machine (56) to determine the health of the electric power train (12).

An air-breathing turbojet engine (301) for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet (302) configured to compress inlet air by one or more shocks, a cooler (303) to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor (304) to compress the air further. A precooler (311) between the inlet (302) and the cooler (303) cools compressed inlet air using compressed air from the turbo-compressor (304). A combustor (305) receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine (306) expands and is driven by combustion products, and a second turbine (307) expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner (309) receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

The invention relates to the technical field of engine recovery, in particular to a turbojet engine ejection recovery system and method in an aircraft fault state, and the system comprises an ejection mechanism which is arranged between a turbojet engine and an aircraft and used for disconnecting the turbojet engine from the aircraft. The turbojet engine is ejected out of the aircraft; the cut-off mechanism is arranged in the aircraft and used for cutting off an oil way and a circuit of the turbojet engine. The protection mechanism comprises a parachute assembly and an air bag assembly, the parachute assembly is connected with the turbojet engine, and a parachute is automatically opened in the landing process of the turbojet engine; the air bag assembly is folded and contracted outside the turbojet engine and automatically expands to wrap the turbojet engine in the landing process. The recycling device is compact in structure and flexible and convenient to use, recycling of the turbojet engine can be efficiently and safely completed when the fixedly-mounted turbojet engine is in an emergency, and the technical requirement for recycling of the turbojet engine in the field is met.

An apparatus includes an engine (102) and an accessory gearbox (106). The engine (102) is positioned on a top side of an aircraft (100). The accessory gearbox (106) is mounted within an airframe (101) of the aircraft (100) remote from the engine (102). The accessory gearbox (106) is drivingly connected to the engine (102).

A jet aircraft having an airflow modification device in its engine duct, the airflow modification device comprising an inlet, an outlet, and a flow path extending from the inlet to the outlet, wherein the device is positioned to receive an incoming airflow and is configured to at least partially modify the airflow within the flow path to improve the ordering of the incoming airflow, the airflow modification device further comprising at least one pitot sensor located at the inlet and integrated with the device, and a pressure line having a first end and a second end, the first end being connected to the pitot sensor and the second end being connected to a transducer, wherein the pressure line is completely enclosed within the device.

Disclosed herein is an aircraft with engines positioned so as to avoid damage to other components of the aircraft. The aircraft may include a blended wing body aircraft. The engines may be positioned aft of a pressure vessel of the aircraft. The engines may be positioned aft of and / or above a fuel store of the aircraft. Such engine positioning may reduce the likelihood of damage to a pressure vessel and / or fuel store in the event of engine failure.

An intentional fluid manipulation apparatus (IFMA) assembly that includes an upstream intentional momentum shedding apparatus (IMSA) configured to impart a first induced velocity to a local free stream flow during a nominal operation requirement. The upstream IMSA creates a streamtube. The IFMA includes a downstream IMSA, with some or all of the downstream IMSA being located in a downstream portion of the streamtube. The downstream IMSA imparts a second induced velocity to the local free stream flow within the streamtube. The second induced velocity at the location of the downstream IMSA has a component in a direction opposite to the direction of the first induced velocity at the location of the downstream IMSA.

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 application discloses a manned aircraft and relates to the technical field of aircraft equipment, and comprises a mounting frame, a first engine, a second engine and a third engine, wherein the first engine, the second engine and the third engine are arranged on the mounting frame, and the output directions of the first engine, the second engine and the third engine are the same; one first engine is arranged at the middle position of the mounting frame, four second engines are arranged at equal intervals in a circle on the periphery of the first engine, and two third engines are symmetrically arranged on the two sides of the four second engines; the axes of the first engine and the two third engines are in the same plane, and the jet ends of the first engine, the second engine and the third engine are respectively provided with an adjusting nozzle, so that the flight attitude control after the manned aircraft is realized.

A craft has: a body including an engine bay; and an engine having an installed condition at least partially in the engine bay. The engine has a heat exchanger having an inlet manifold; and in the installed condition an air flowpath passes through the inlet manifold to the heat exchanger and exits the heat exchanger to the engine bay.

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.

PROPULSION UNIT, AIRCRAFT AND IN-FLIGHT OPERATION The present invention relates to a propulsion unit (100) for a fixed-wing aircraft, the propulsion unit comprising a nacelle (112) housing at least two non-coaxial fans (14), said nacelle comprising an upper wall (116) having a span overhanging the two fans, notable in that the upper wall comprises a nozzle pivotable (116.2) about an axis (116.3) parallel to the span of the upper wall. The invention also relates to a method of operating an aircraft comprising a climb phase during which the nozzle is radially inclined outwards, and potentially a cruise flight phase during which the nozzle remains stationary, the speed of the aircraft's fans varies, the inclination of the nozzle relative to the upper wall varying proportionally to the speed of the fans. (Figure to be published with the abbreviation: Figure 4)

This aircraft hybrid motive power source system is installed in an aircraft and comprises: a gas turbine engine (20) that includes a low-pressure shaft (27) and a high-pressure shaft (28) as rotating shafts; a first electric generator (31) that is drivably connected to the high-pressure shaft (28); a second electric generator (32) that is drivably connected to the low-pressure shaft (27); a fuel cell (40) to which fuel gas and oxidant gas are supplied; a fuel gas generation unit (50) that generates fuel gas from a raw material through heating using exhaust gas of the gas turbine engine (20); and a control unit (60) that, on the basis of information indicating the running state of the gas turbine engine (20) and / or demand for electric power in the aircraft, supplies, to the first electric generator (31) or the second electric generator 32, electric power of the fuel cell (40) obtained through the supply of fuel gas.

The present invention relates to a system (30) for controlling actuators (TS) comprising a plurality of redundant servo units (FC-1, FC-2, FC-3) generating values ​​of the same magnitude (Δ1, Δ2, Δ3) from the same setpoint (Gsp) and the same measurement data (Gs1, Gs2, Gs3), the redundant servo units (FC-1, FC-2, FC-3) further communicating with each other (B) so that each redundant servo unit (FC-1, FC-2, FC-3) has access by reading the most recent values ​​of the variables (Δ1, Δ2, Δ3) generated by the redundant servo unit and implements a method for correcting each generated value by combining them. Thus, the present invention prevents discrepancies between these variables (Δ1, Δ2, Δ3). Such a system (30) may be equipped in a vehicle having a driver, passengers, and / or a load consisting of goods or cargo, and having one or more actuators controlled in this manner.