Abstract: In a flying object, a PCU has a plurality of operation modes in which an engine and/or a motor generator is used as a driving source for a pusher propeller. In accordance with the state of the flying object, the PCU controls the engine, a clutch, and the motor generator in one of the operation modes.

Abstract: An aircraft power system comprises a turbocharger, the turbocharger including a compressor for supplying combustion air to an internal combustion engine, a turbine operatively connected to an internal combustion engine to receive an exhaust flow from the internal combustion engine and convert energy of the exhaust flow into rotational power and, a turbo shaft operatively connecting the turbine to the compressor to transfer at least some of the rotational power to the compressor. A generator is operatively connected to the turbo shaft to receive at least some of the rotational power from the turbo shaft for generating electrical power. At least one electrically powered air-mover is electrically connected to the generator to receive at least some of the electrical power to produce thrust.

Abstract: An air mobility vehicle may include air flaps located under a mounting position of each of rotary wings and rotatably mounted inside openings to guide a flow direction of air flowing to a region under each of the rotary wings or air flowing above the openings to inside of the air mobility vehicle, an actuator coupled to the air flaps and configured to rotate the air flaps to direct the air having passed through the air flaps to a motor, an inverter, or the motor and the inverter of each of the rotary wings, or batteries, and a controller electrically connected to the actuator and configured to control a flow of the air having passed through the air flaps by controlling the actuator depending on a driving state of the air mobility vehicle or temperatures of the motor and the inverter of each of the rotary wings or a temperature of the batteries.

Abstract: The invention relates to an aircraft propulsion system (100) intended for being built into the rear of an aircraft fuselage, the propulsion system comprising at least two gas generators (102a, 102b) supplying a power turbine (104) having two counter-rotating turbine rotors (104a, 104b) for driving two fans (112a, 12b), and separate air inlets (106a, 106b) for supplying each gas generator, characterised in that it comprises an electrical drive device (140) configured to rotate at least one of the turbine rotors, at least one electrical generator (142a, 142b) configured to transform part of the energy of the flow from the gas generators into electrical power and an electric motor (146) supplied by said electrical generator and capable of rotating at least one of the turbine rotors, said electrical generator being installed on one of said gas generators, and in that said turbine rotor is capable of being rotated simultaneously by a flow from said gas generators and by the electrical drive device.

Abstract: A general mounting method for ducted fan propulsors is disclosed. This mounting method uses extremely slender stators that connect the duct ring to the propulsor which is mounted in the middle and drives the rotor, fan or propeller. The slender stators take the form of spokes and as such are so slender that the midspan stresses within the spokes are dominated by axial tension loads rather than the shear loads experienced by conventional stators. The spokes may have an aerodynamic shape and damping methods may be used to retard spoke vibrations and transmission of engine vibrations to the duct and force. The duct itself is also stiffened by the spoke arrangement, thereby reducing low frequency vibration modes.

Abstract: A propulsion device includes: a duct in which a flow path extending in a direction of an axis; a fan which is provided with a) an outer peripheral ring formed in an annular shape surrounding the axis and installed to be relatively rotatable around the axis with respect to the duct, b) a plurality of fan blades arranged at intervals in a circumferential direction such that each blade is extended from the outer peripheral ring toward the inside of the flow path, and c) an inner peripheral ring formed in an annular shape being connected radially inner end portions of the plurality of fan blades and in which an air flow passage is formed so that air flows therethrough in the direction of the axis; and a motor which drives the fan to rotate around the axis.

Abstract: An aircraft propulsion plant including an electric motor having a rotor and a stator mechanically linked to a base which can be mounted at the rear of an aircraft fuselage, a fan rotated by the rotor, a set of fixed blades located downstream of the fan, and a nacelle comprising an outer casing and a fan casing surrounding the fan and the set of fixed blades. The nacelle is mechanically linked to the base through the set of fixed blades. This configuration enables a cooling circuit to be formed for enabling the heat produced by the electric motor at the location of the stator to be evacuated towards the fixed blades and the nacelle where it is dissipated. Furthermore, this heat may be used for the de-icing of the nacelle lip.

Abstract: A propulsion system for omnidirectional maneuverability and efficient forward flight of an airship. The propulsion system includes only fixed, unidirectional engines (17, 19, 20). Thrust vectors of the fixed engines (19, 20) are oriented in a way that their speeds can be chosen such that all forces acting on the airship (i.e., engine thrusts, gravity, buoyancy, wind and potentially others) together result in the desired motion. The engines may be four ducted fans (17) at the bow of the aircraft and four ducted fans (19) at the stern of the aircraft. The thrust vectors of the engines can be decomposed into three vectors of equal length that are each parallel to one of the three axes of a Cartesian coordinate system. Efficient forward flight is achieved by an additional engine (20) at the stern of the airship.

Abstract: An aircraft is provided including a fuselage that extends along a longitudinal direction between a forward end and an aft end. A boundary layer ingestion fan is mounted to the fuselage at the aft end and is configured for ingesting boundary layer airflow off the surface of the fuselage. The fuselage defines a profile proximate the boundary layer ingestion fan that is optimized for ingesting a maximum amount of boundary layer air and improving propulsive efficiency of the aircraft. More specifically, the fuselage defines a cross sectional profile upstream of the boundary layer ingestion fan that has more cross sectional area in a top half relative to a bottom half as defined relative to a centerline of the boundary layer ingestion fan.

Abstract: An aircraft defining a longitudinal centerline and extending between a forward end and an aft end is provided. The aircraft includes a fuselage extending longitudinally between the forward end of the aircraft and the aft end of the aircraft; a primary wing assembly extending laterally outwardly with respect to the longitudinal centerline from a portion of the fuselage; a first engine mounted to the primary wing assembly; and a first engine wing assembly extending outward from the first engine.

Abstract: A modular wing in ground effect vehicle has a fuselage which receives interchangeable cockpit and component modules and may accept a variety of snap-on wing styles and control surfaces for various flying conditions and pilot skill levels. When depleted or due for repairs, modules containing stored energy components or machinery subject to wear or periodic maintenance may be exchanged for fresh units so that a vehicle requiring such an exchange may be returned to service quickly and conveniently.

Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan defining a fan tip radius at the fan face; wherein a downstream blockage ratio is defined as: the ? ? turbine ? ? diameter ? ? at ? ? an ? ? axial ? ? location of ? ? the ? ? lowest ? ? pressure ? ? rotor ? ? stage a ? ? distance ? ? f ? rom ? ? a ? ? ground ? ? plane ? ? to ? ? the ? ? wing ? and wherein an engine blockage ratio of: ( 2 × the fan tip radius/the engine length ) the ? ? downstream ? ? blockage ? ? ratio is in the range from 2.5 to 4.

Abstract: An arrangement for use with a propulsion system for a supersonic aircraft includes a center body configured for coupling to an inlet and to support a boundary layer formed when the supersonic aircraft is flown at a predetermined altitude supersonic speed. The arrangement further includes a first vortex generator disposed on the center body. The first vortex generator extends a first height above the center body. The arrangement still further includes a second vortex generator disposed on the center body. The second vortex generator extends a second height above the center body, the second height being greater than the first height. The first height and the second height are greater than approximately seventy-five percent of a thickness of the boundary layer proximate a location of the first vortex generator and the second vortex generator, respectively, when the aircraft if flown at the predetermined altitude and the predetermined speed.

Abstract: An airplane propelled by a turbine engine having at least one fan, the turbine engine being integrated in the rear of a fuselage of the airplane, extending it rearwards, the airplane further including at least one acoustic baffle forming panel connected to the fuselage of the airplane and arranged below the turbine engine.

Abstract: A propulsion system for an aircraft includes an engine (e.g., piston engine, or turbine) having an axis made to be nonparallel with the longitudinal axis of the aircraft. This is enabled using an electrical, hydraulic, or other system to transfer energy generated by the engine (e.g., via electrical wiring, fluid conduits, etc.) to remotely power a motor used to drive a thrust-creating device (e.g., propeller or ducted fan). That the engine is able to be freely oriented allows for it being positioned without regard to any mechanical restraints existing in conventional arrangements.

Abstract: A gas turbine engine includes a compressor section and a turbine section. The turbine section includes a drive turbine and is located downstream of the compressor section. The gas turbine engine also includes a fan mechanically coupled to and rotatable with the drive turbine such that the fan is rotatable by the drive turbine at the same rotational speed as the drive turbine, the fan defining a fan pressure ratio and including a plurality of fan blades, each fan blade defining a fan tip speed. During operation of the gas turbine engine at a rated speed, the fan pressure ratio of the fan is less than 1.5 and the fan tip speed of each of the fan blades is greater than 1,250 feet per second.

Abstract: The present disclosure is directed to a fan section mounted to an aircraft tail section. The fan section defines a radial direction, an axial direction and a circumferential direction. The fan section includes a shaft extending generally along a longitudinal axis, and a plurality of fan blades rotatable with the shaft about the longitudinal axis. Each fan blade defines an outer end along the radial direction. One or more of the plurality of fan blades translates a respective outer end from a first radial position to a second radial position.

Abstract: An aircraft has an internal combustion engine and a wing including a hollow structural member. The aircraft has an electrical network including: an electrical motor configured to drive an aircraft propulsor; at least one conductor configured to electrically couple the electrical motor and an electrical storage device; wherein the electrical conductor is formed of the hollow structural member of the aircraft wing.

Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades, each of the plurality of fan blades defining a fan blade span along a radial direction. The gas turbine engine further includes an outer nacelle surrounding the plurality of fan blades and a plurality of part-span inlet guide varies attached to the outer nacelle at a location forward of the plurality of fan blades along an axial direction. Each of the plurality of inlet guide vanes defines an IGV span along the radial direction, the IGV span being at least about five percent of the fan blade span and up to about fifty-five percent of the fan blade span.

Abstract: The present invention includes a distributed propulsion system for a craft that comprises a frame, a plurality of hydraulic or electric motors disposed within or attached to the frame in a distributed configuration; a propeller operably connected to each of the hydraulic or electric motors, a source of hydraulic or electric power disposed within or attached to the frame and coupled to each of the disposed within or attached to the frame, wherein the source of hydraulic or electric power provides sufficient energy density for the craft to attain and maintain operations of the craft, a controller coupled to each of the hydraulic or electric motors, and one or more processors communicably coupled to each controller that control an operation and speed of the plurality of hydraulic or electric motors.

Abstract: An aircraft has an internal combustion engine and a wing including a hollow structural member. The aircraft has an electrical network including: at least one alternating current electrical generator configured to be driven by the internal combustion engine; an electrical motor configured to drive an aircraft propulsor; at least one conductor configured to electrically couple the electrical motor and the electrical generator; wherein the electrical conductor is formed of the hollow structural member of the aircraft wing.

Abstract: A drive arrangement for an aircraft comprises a pair of propulsor units each having a fan and a fan shaft for driving the fan. A core engine has a turbine driving a core engine shaft. A mechanical connection connects the core engine shaft to drive the fan shafts for each of the propulsor units. An aircraft also has such an arrangement.

Abstract: A turbofan engine including an axially extending inlet wall surrounding an inlet flow path. A radial distance between the inlet wall and the inner wall adjacent the fan defines a downstream height of the inlet flow path. A plurality of vanes are circumferentially spaced around the inlet, each of the vanes extending radially inwardly from the inlet wall, a maximum radial distance between a tip of each of the vanes and the inlet wall defining a maximum height of the vane. The maximum height of the vane is at most 50% of the downstream height of the flow path. In another embodiment, the maximum height of the vane is at most 50% of the maximum fan blade span. A method of reducing a relative Mach number at fan blade tips is also discussed.

Abstract: A propulsion engine comprising at least a first propulsion unit including a first fan encased by a geared ring and a gas turbine engine driving a first shaft connected to the first fan, at least a second propulsion unit including a second fan encased by a geared ring connected to a second shaft operatively coupled to an electric machine and at least an electric storage device connected to the electric machine. The geared rings are configured to transmit torque between the fans so that they can rotate in conjunction (directly or through an intermediate gear) when they are driven by at least one of said first and second shafts. The propulsion engine is arranged for controlling the torque to be supplied to the assembly of the first and second fans by the gas turbine engine and/or by the electric machine acting as a motor or as a generator.

Abstract: Propulsion unit for an aircraft, comprising a turbine, at least one propeller offset with respect to the turbine, and a power transmission disposed between the turbine and the propeller, the transmission comprising in series two constant velocity joints with a slide connection.

Abstract: An aircraft (10) comprising a fuselage (12), the fuselage (10) comprising at least first and second inwardly tapering portions (20, 22) extending in a generally rearward direction from an aft end of the fuselage (12), each inwardly tapering portion (20, 22) comprising a boundary layer ingesting propulsor (36) mounted thereon.

Abstract: The invention relates to an architecture of a propulsion system of a multi-engine helicopter, comprising turboshaft engines (1, 2) that are connected to a power transmission gearbox (3), and comprising a low DC voltage onboard network (7) for supplying helicopter equipment during flight, characterized in that it comprises: a hybrid turboshaft engine (1) that is capable of operating in at least one standby mode during a stable flight of the helicopter; an electrotechnical pack (20) for quickly restarting said hybrid turboshaft engine in order to bring said engine out of said standby mode and to reach a mode in which it provides mechanical power, said restart pack (20) being connected to said onboard network (7); and at least two sources (4, 16, 18) of electrical power for said onboard network (7).

Abstract: An aircraft fuselage frame including a central element adapted to be located within the perimeter of the fuselage, and two lateral extensions projecting outside the perimeter of the fuselage from both sides of the central element that are a portion of a longitudinal structure of a lifting surface. The central element and the two lateral extensions are configured as an integrated piece.

Abstract: A turbofan engine including an axially extending inlet wall surrounding an inlet flow path. A radial distance between the inlet wall and the inner wall adjacent the fan defines a downstream height of the inlet flow path. A plurality of vanes are circumferentially spaced around the inlet, each of the vanes extending radially inwardly from the inlet wall, a maximum radial distance between a tip of each of the vanes and the inlet wall defining a maximum height of the vane. The maximum height of the vane is at most 50% of the downstream height of the flow path. In another embodiment, the maximum height of the vane is at most 50% of the maximum fan blade span. A method of reducing a relative Mach number at fan blade tips is also discussed.

Abstract: An aerial vehicle includes a hybrid power generation system comprising an engine; a generator mechanically coupled to the engine; and a propulsion system comprising an electric motor electrically coupled to the generator and a rotational mechanism coupled to the electric motor.

Abstract: Aeroshells and methods for manufacturing aeroshells are provided. In this regard, a representative aeroshell for a missile formed of a long fiber thermoplastic composite exhibits a wall thickness of no greater than approximately 0.070?.

Abstract: An aircraft including a fuselage and an aft engine is provided. The fuselage extends from a forward end of the aircraft towards an aft end of the aircraft. The aft engine is mounted to the fuselage proximate the aft end of the aircraft. The aft engine includes a fan rotatable about a central axis of the aft engine, the fan including a plurality of fan blades. The aft engine also includes a nacelle surrounding the plurality of fan blades and defining an inlet. The inlet defines a non-axis symmetric shape with respect to the central axis of the aft engine to, e.g., allow for a maximum amount of airflow into the aft engine.

Abstract: An aircraft including a fuselage having a forward portion and an aft portion with a propulsion system mounted within the aft portion of the fuselage. A burst zone is defined that extends outward from the propulsion system. The aircraft includes a box wing extending from the aft portion of the fuselage to a forward portion of the fuselage that is disposed outside of the burst zone.

Abstract: An air vehicle may include a body having forwardly-swept fixed wings or unswept fixed wings. The body may have a longitudinal axis continuously oriented parallel to a forward flight direction when the air vehicle is in flight. The air vehicle may include at least one of the following components being mounted on opposite sides of the body: longitudinally offset engine nacelles, asymmetrically lengthened engine nacelles, and longitudinally offset protruding aerodynamic surfaces comprising at least one of stabilizers, canards, empennage control surfaces, external stores, and high aspect ratio wings.

Abstract: An airframe structural element, such as an aircraft rib, is disclosed that comprises an integral pipe connector element for connecting pipes through the structural element. The pipe connection means provided by the pipe connector element may be aligned on a common axis with a pipe connection means provided by an adjacent airframe structural element.

Abstract: A method for operating an aircraft that is driven forward in the air with a propulsion system in the form of two engines of the jet or propeller type one on each side of the fuselage is disclosed. During forward movement of the plane, the two engines move in a mutual synchronous relationship relative to the longitudinal axis of the plane. The propulsion system is used for aircraft as well as for ships.

Abstract: An apparatus is described comprising: at least one processor; and memory storing instructions that, when executed by the at least one processor, cause the apparatus to: regulate an operation of a motor based at least in part on a control imposed on an output torque of the motor and a rate of change of a speed of the motor.

Abstract: An aircraft nacelle includes at its outside wall a cowl moveable relative to the nacelle so as to block or unblock an opening. The cowl includes an articulation and locking/unlocking elements distant from the upstream edge of the cowl that include elements for limiting the appearance of scooping phenomena. The limiting elements include at least one lock integral with the rest of the nacelle or respectively the cowl that can occupy a locked state in which part of the lock interferes with the cowl or respectively the rest of the nacelle and prevents a deformation that includes a radial component of the cowl, and another free state in which the part no longer interferes with the cowl or respectively the rest of the nacelle and allows the opening movement of the cowl, with the switch from one state to the next being generated by a gas pressure variation.

Abstract: An air vehicle having a bilaterally asymmetric configuration for reducing wave drag may include a body having a longitudinal axis. The air vehicle may further include longitudinally offset engine nacelles, asymmetrically lengthened engine nacelles, and/or longitudinally offset protruding aerodynamic surfaces for reducing wave drag.

Abstract: A structural rod has a primary rod that has an axial cavity and a secondary rod inside the axial cavity so that the secondary rod is able to move inside the axial cavity along the direction of the longitudinal axis of the rod. The primary rod has a bearing for fastening the structural rod at the first end in a rigid manner and is not fastened rigidly to the structure at the second end. Conversely, the secondary rod has a bearing for fastening the structural rod at the second end in a rigid manner and is not fastened rigidly to the structure at the first end. The primary and secondary rods are joined through one or more intermediate linkage elements made of a deformable and elastic material.

Abstract: A tail section for an aerospace vehicle is provided. The tail section comprises a rudder which is movable about an axis to generate a yawing moment on the aerospace vehicle. The tail section further comprises a thruster having, in flow series, an air intake, an electrically powered device for accelerating the air received through the intake, and an air outlet which directs the accelerated air to increase the yawing moment generated by the rudder.

Abstract: An aircraft engine assembly including a mounting pylon including a rigid structure fitted with a central box, together with a forward extension and a rear extension each including an attached upper organ supported on an upper surface of the central box, and an attached lower organ supported on a lower surface of the front end. The forward and rear extensions support at least a part of the first fasteners interposed between the rigid structure and a turbine engine.

Abstract: A turbofan engine control system and method includes a core nacelle housing (12), a compressor and a turbine. A turbofan is arranged upstream from the core nacelle and is surrounded by a fan nacelle (34). A bypass flow path (39) is arranged downstream from the turbofan between the core and fan nacelles. The bypass flow path includes a nozzle exit area (40). A controller (50) detects at least one of a take-off condition and a landing condition. The controller changes effectively the nozzle exit area to achieve a thrust vector in response to the take-off and landing conditions.

Abstract: An aircraft may have a fuselage, a left wing extending from the fuselage, a right wing extending from the fuselage, a tail section extending from a rear portion of the fuselage, and a first engine and a second engine operably connected by a common driveshaft, wherein the first and second engines are configured for freewheeling such that if one of the first and second engines loses power the other of the first and second engines continues to power the aircraft.

Abstract: An aircraft is provided includes at least one drive flow passage, which runs from an air inlet directed forward on the body surface via a jet engine through the body to a jet nozzle that opens towards the rear on the body surface. At least a part of the jet engine is arranged upstream of the air inlet seen in the flight direction of the aircraft and the drive flow passage has curvature sections embodied and arranged for this in a suitable manner.

Abstract: The present invention relates to a power plant (10) having a piston engine (30) driving a conventional gearbox (20) dimensioned for a turbine engine, said piston engine (30) comprising a first row (31) of pistons (33) presenting an angle (?) relative to a second row (32) of pistons (33). The power plant includes a flywheel (50) and a torsion shaft (40) set into motion by said pistons (33), said torsion shaft (40) being arranged between the first and second rows (31, 32) of pistons.

Abstract: An aircraft wing turbine engine is designed to increase the ground clearance for larger modernized engines mounted under the wings of an airfoil. To this end, the aircraft wing turbine engine includes a hot flow generator with a turbine engine rotating about a first central axis, a cold flow blower rotating about a second central axis, and a pod surrounding the hot flow generator and cold flow blower. The first and second central axes are offset and non-collinear with each other such that the cold flow blower can be moved farther away from the ground to increase the ground clearance of the turbine engine.

Abstract: An engine mounting arrangement includes a propulsor mounted to an aircraft wing, and an engine core aerodynamically connected to the propulsor and positioned rearward of the propulsor.

Abstract: An aircraft pylon support system for mounting and supporting aircraft propulsion systems to a fuselage section of an aircraft that includes an internal, central support box disposed within and extending across the interior of an aircraft fuselage, and external support boxes extending laterally into pylons disposed at both sides of the aircraft, each of the internal and external support boxes being made of a composite material and structured as multi-spar boxes containing lateral spars and at least one central spar, wherein the central and lateral support boxes are connected together to provide a continuous interface therebetween with a full transfer of loads therebetween and between any interrupted, intermediate frame members.

Abstract: The invention relates to multimode aircraft operated at supersonic and subsonic flight speeds in a wide range of flight altitudes. The invention is particularly applicable in multimode super-maneuverable aircraft capable of cruising at supersonic speed and having low radar signature. The technical effect to be achieved by the invention is to provide an aircraft which has low radar signature, is super-maneuverable at high angles of attack, has high aerodynamic efficiency at supersonic speeds while keeping high aerodynamic efficiency at subsonic modes, and can accommodate bulky cargo in its internal compartments. The integral aerodynamic configuration aircraft comprises fuselage (1) with strake (2), wing with panels (3) smoothly blended into fuselage (1), all-moving horizontal tail (AMHT) (4), and all-moving vertical tail (AMVT) (5). Mid-fuselage is flattened and formed longitudinally by a set of aerodynamic profiles.