606results about "Jet type power plants" patented technology

An aircraft defining a longitudinal centerline and extending between a forward end and an aft end is provided. The aircraft comprises a fuselage extending between the forward end of the aircraft and the aft end of the aircraft; a wing assembly extending laterally outwardly with respect to the longitudinal centerline from a portion of the fuselage; an unducted turbofan engine including an unducted fan that defines a plurality of fan wakes; and a pylon fairing that connects the unducted turbofan engine to the wing assembly, the pylon fairing defining a pylon fairing dihedral, wherein the pylon fairing dihedral is in a direction opposite the incoming fan flow swirl and associated sheared profiles of fan wakes upon arrival at the pylon leading edge.

An aircraft defining a longitudinal centerline and extending between a forward end and an aft end is provided. The aircraft comprises a fuselage extending between the forward end of the aircraft and the aft end of the aircraft; a wing assembly extending laterally outwardly with respect to the longitudinal centerline from a portion of the fuselage; an unducted turbofan engine including an unducted fan defining a fan diameter; and a pylon fairing that connects the unducted turbofan engine to the wing assembly, the pylon fairing defining a leading edge and including a first inflection point along the leading edge, the first inflection point defining a first radius of curvature that is less than 5 times the fan diameter and greater than 0.1 times the fan diameter.

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.

This invention relates to a method for controlling a motor assembly (100). This motor assembly (100) comprises at least a first electric machine (300) and a gas turbine engine (200). The gas turbine engine (200) comprises a low-pressure shaft (210) and a high-pressure shaft (220). The electric machine (300) is coupled to the low-pressure shaft (210), and the control method comprises a step in which a take-off of mechanical work is ordered from the first electric machine (300) to brake a rotation of the low-pressure shaft (210) in response to an activation of a thrust reverser (281) of the gas turbine engine (200) and / or a disturbance of the air flow in a transverse plane at an air intake of the gas turbine engine (200). The invention also relates to a control unit (500) suitable for carrying out this method, a motor assembly (100) incorporating this control unit (100), the electric machine (300) and the gas turbine engine (200), and a computer program to carry out this method.

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.

Aerodynamic lifting structures, such as aircraft wings, having embedded engines and associated methods and systems are disclosed herein. A wing assemblies configured in accordance with embodiments of the present technology can include, for example, an upper wing portion, a lower wing portion, and a plurality of independent ducts positioned between the upper wing portion and the lower wing portion. Each duct can extend between a corresponding inlet positioned toward a leading portion of the wing assembly and a corresponding outlet positioned toward a trailing portion of the wing assembly. The wing assembly can further include a plurality of fans and a plurality of electric motors operably coupled to the plurality of fans. The fans and electric motors are positioned in the corresponding individual ducts and the fan is rotatable to propel fluid received in the inlet through the duct to create lift.

The invention discloses a tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system.The tilt-rotor vertical take-off and landing aircraft and the vertical take-off and landing balance system.The tilt-rotor vertical take-off and landing aircraft comprises an aircraft body, wings are symmetrically arranged on the two sides of the aircraft body, wingtip small wings are arranged at the tail ends of the wings, and the wingtip small wings are arranged on the aircraft body. A pair of tilting mechanisms symmetrically arranged relative to the fuselage is arranged at the front end of the fuselage, each tilting mechanism comprises a front propeller, the front propellers are connected with the fuselage through tilting motor structures, and a micro turbojet engine is arranged at the rear end of the fuselage and connected with the fuselage through a fixing frame. The vertical take-off and landing balance system comprises a hardware part and a software part. Smooth transition between a vertical take-off and landing mode and a horizontal flight mode is achieved through optimized mode switching control logic.

A system for an aircraft includes a fan rotor, an engine core, a fan case, a core case, a bifurcation, an inner electrical system, an outer electrical system and multiple electrical lines electrically coupling the inner electrical system to the outer electrical system. The engine core is configured to drive rotation of the fan rotor about an axis. The fan case houses the fan rotor. The core case houses the engine core. The bifurcation radially crosses a bypass flowpath. The bypass flowpath is downstream of the fan rotor and located radially between the fan case and the core case. The inner electrical system is arranged with the engine core. The outer electrical system is mounted to the fan case. The electrical lines extend between and electrically couple the inner electrical system to the outer electrical system. The electrical lines run in parallel along one another radially through the bifurcation.

The invention provides an aircraft framework and an aircraft, the aircraft framework is used for the aircraft flying in the atmosphere, the aircraft framework comprises a cage serving as a framework main body, the cage bears the self weight and the flight load of the whole aircraft, at least two first turbojet engines can be contained in the cage, and the first turbojet engines are arranged in the cage; the cage is open at least in the lower portion and used for allowing the jet flow of the first turbojet engine to pass through. Through the arrangement of the cage which is a skeleton main body, a power source device of the aircraft is collected and gathered, so that the miniaturization of the whole aircraft is realized.

A propulsion assembly is disclosed including a pylon, a turbomachine having an engine and a fan surrounded by a fan casing, a nacelle surrounding the engine and the fan casing, and including a load bearer positioned at the upper part of the nacelle, a forward fan mount between a frontal zone of the fan casing and a frontal zone of the load bearer, a rear pylon mount between the pylon and the load bearer, and a positioning assembly providing positioning between the fan casing and the load bearer. The propulsion assembly is configured to control the separations between the cowls mounted in a moveable manner on the load bearer and a nose cowl.

The invention provides a personal aircraft chassis and a personal aircraft. The chassis comprises a main framework and a chassis body, the personal aircraft comprises a main framework and a thrust array assembly, the thrust array assembly is fixed to the main framework, thrust of the thrust array assembly is achieved through a plurality of first turbojet engines, the geometric position relation between the thrust action point and the thrust direction of the thrust array assembly and the main framework is kept fixed, and when the personal aircraft stops on the ground, the thrust array assembly is fixed to the main framework. The thrust direction of the thrust array assembly is perpendicular to the ground. The vector nozzle assemblies are fixed to the main framework, thrust of the vector nozzle assemblies is achieved through the second turbojet engines, the central axes of the second turbojet engines are perpendicular to the central axis of the first turbojet engine, and vector nozzles of the vector nozzle assemblies at least have one degree of freedom relative to the main framework. And high-dynamic, high-fault-tolerance and reconfigurable flight control which cannot be realized by utilizing pure thrust difference control can be realized.

A fitting to mount an engine to a wing of an aircraft. The fitting includes a body with a first end and a second end. A first pocket is positioned at the first end with the first pocket sized to receive the forward engine mount fitting. A second pocket is positioned at the second end with the second pocket sized to receive a strut of the wing of the aircraft. The body includes a first section and a second section that are connected together.

An electric power system (400) for a vehicle includes at least one electric machine (321), one or more power rectifiers, and a plurality of DC channels. The at least one electric machine (321) includes a plurality of tooth-wound multi-phase windings (311) that are substantially magnetically decoupled, and the at least one electric machine (321) is mechanically balanced even if one of the plurality of windings (311) is de-energized. The one or more power rectifiers are configured to produce rectified power from the power generated by the at least one electric machine (321). The plurality of DC channels are formed after the at least one power rectifier and are configured to provide DC power to one or more loads within a vehicle.

An aircraft assembly is disclosed having a wing and an engine mounting pylon. An aft end of the engine mounting pylon is connected to the wing by a spigot and at least one fastener. The aircraft assembly is configured such that, during operation of the aircraft assembly on an aircraft, the spigot transfers only lateral load between the engine mounting pylon and the wing and the at least one fastener transfers only vertical load between the engine mounting pylon and the wing.

An aircraft hybrid motive power source system includes: a gas turbine engine mounted on an aircraft and including a low-pressure shaft and a high-pressure shaft as rotary shafts; a first motor generator drivingly connected to the high-pressure shaft; a second motor generator drivingly connected to the low-pressure shaft; a fuel cell to which a fuel gas and an oxidant gas are supplied; a fuel gas generator configured to generate the fuel gas from a raw material by heating using an exhaust gas from the gas turbine engine; and a controller configured to supply an electric power of the fuel cell obtained by a supply of the fuel gas to one of the first motor generator and the second motor generator, based on at least one of information indicating an operating status of the gas turbine engine and an electric power demand of the aircraft.

A support assembly for supporting a main accessory gearbox of an aircraft turbine engine, the accessory gearbox including gears and supporting at least one item of equipment driven by the gears. The support assembly can include a structure for connecting and supporting the turbine engine to a pylon of the aircraft including: an intermediate axial portion for attachment to the pylon, having an upper end that defines a linking interface with the pylon, the intermediate axial portion bearing suspension rods, which are intended to be connected to the turbine engine; a front axial portion extending forward of the intermediate portion and having at least one suspension member configured to be connected to the turbine engine; and a rear axial portion extending rearwards of the intermediate portion and supporting the main accessory gearbox.

A structure for linking and supporting a turbine engine on an aircraft pylon, the structure having a first longitudinal axis intended to extend parallel to a second longitudinal axis of the pylon, the structure including a first axial portion for fastening to the pylon, the first portion having an upper end that defines a substantially horizontal plane of interface with the pylon, and a lower end bearing suspension rods for suspending the turbine engine, which extend in a rear vertical plane; and a second axial portion that is intended to extend forward of the pylon and includes at least one suspension member of the turbine engine in a front vertical plane.

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.

An assembly including a wing with a pressure-side panel and a suction-side panel through which two bores pass, an engine pylon including a blade with three bores and two fastening lugs, each of which has a sixth bore, two first shackles fastened between the blade and the suction-side panel, two second shackles fastened between the blade and the suction-side panel and two fittings, which are fastened beneath the pressure-side panel and each have a female clevis for housing a fastening lug therein.

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 method of operating an aircraft engine coupled to a wing of an aircraft including: setting a pitch of a plurality of rotor blades of a rotor assembly of the aircraft engine at non-uniform pitch angles along a circumferential direction of the aircraft engine such that the plurality of rotor blades define a first pitch at a first position and a second pitch at a second position, wherein the second position is 180 degrees offset from the first position, and wherein the first pitch is different from the second pitch.

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.

A strut assembly for coupling an engine to a wing of an aircraft, the strut assembly comprising a mid truss assembly comprising a fore portion and an aft portion; an engine mount member connected to the fore portion of the mid truss assembly; an aft truss assembly comprising a fore portion and an aft portion, the fore portion of the aft truss assembly being connected to the aft portion of the mid truss assembly; an aft strut bulkhead connected to the aft portion of the aft truss assembly; and a wing mounting feature operatively connected to, and located aft of, the aft strut bulkhead.

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.