Abstract: A system for generating accessory power from a gas turbine engine is provided by the present invention. The system includes an electronic control device for monitoring at least one parameter which provides information about an incipient change in power demand, a control valve operated by the control device for supplying bleed air from the engine during a transient state in response to the at least one monitored parameter, and a pneumatically operated device for receiving the bleed air and for generating power to operate equipment onboard an aircraft. The pneumatically operated device may be an air turbine or a pneumatically integrated generator.

Abstract: A hybrid power plant (5) for an aircraft (1) comprises at least: a hybrid drive system (37) having a main on-board electricity network (16) and an auxiliary electricity network (34); and a selective adaptation interface (38) arranged to enable electrical energy to be exchanged selectively between the main and auxiliary electricity networks (16; 34). At least one engine and a hybrid drive auxiliary electrical machine (7, 31) are mechanically connected to a transmission (8); said machine (7) being electrically connected to at least one auxiliary electrical bus (36) in parallel with at least one auxiliary device for delivering electric charge.

Abstract: Vehicles, systems, and methods are disclosed for providing and directing first power, such as vehicle generator power, and alternate sources of power. In a particular embodiment, a vehicle includes a power distribution grid that includes a plurality of power sources and a plurality of distributions buses configured to distribute power from the plurality of power sources. The plurality of power sources include an engine-driven power source is configured to provide first power where the first power has first power characteristics. The plurality of power sources also includes a plurality of engine-independent power sources including a first alternate power source configured to provide first alternate power. The first alternate power has first alternate power characteristics that are different than the first power characteristics. The plurality of engine-independent power sources also includes a second alternate power source configured to provide second alternate power.

Abstract: An air vehicle incorporating a hybrid propulsion system. The system includes a gas turbine engine as a first motive power source, and one or more battery packs as a second motive power source. Through selective coupling to a DC electric motor that can in turn be connected to a bladed rotor or other lift-producing device, the motive sources provide differing ways in which an aircraft can operate. In one example, the gas turbine engine can provide operation for a majority of the flight envelope of the aircraft, while the battery packs can provide operation during such times when gas turbine-based motive power is unavailable or particularly disadvantageous. In another example, both sources of motive power may be decoupled from the bladed rotor such that the vehicle can operate as an autogyro.

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: A gearbox assembly includes a first face gear driveable about a face gear axis in a first direction by at least one first pinion gear and a second face gear driveable about the face gear axis in a second direction opposite the first direction by at least one second pinion gear. A thrust bearing is located between the first face gear and the second face gear. The first face gear is configured to drive rotation of a first shaft in the first direction and the second face gear is configured to drive rotation of a second shaft in the second direction. Further disclosed is a power train for a rotary wing aircraft having a gearbox including a first face gear rotatable in a first direction and a second face gear rotatable in a second direction opposite the first direction.

Abstract: A gyrocopter includes a motor, a rotor head and a torque transmission device for transmitting torque of the motor to the rotor head. The torque transmission device includes a pneumatic coupling.

Abstract: A gas turbine engine includes a spool along an engine centerline axis which drives a gear train, said spool includes a low pressure compressor with four (4) stages.

Abstract: The present application relates to a regenerative or hybrid drive system for an aircraft. The drive system includes the use of a power augmentation device to selectively engage the drive system of an aircraft to store power. The power in the power augmentation device being configured to release a power burst to systems within the aircraft to perform critical flight maneuvers.

Abstract: A rotary-wing aircraft includes a retractable port movable between a closed position and an open position to selectively control flow of combustion gases into a bypass passage to change a power distribution between a rotor system and a secondary propulsion system.

Abstract: An assembly includes an output shaft, an input shaft with an internal reservoir, and a clutch. The clutch is adapted to selectively engage and couple the input shaft to the output shaft. The clutch has a pressure chamber that communicates with the reservoir of the input shaft. The reservoir has a sufficient volume to provide the pressure chamber with a fluid pressure adequate to maintain engagement of the clutch during an interruption of a fluid supply to the input shaft and clutch.

Abstract: An aircraft rotor constant-velocity drive system having a differential mechanism that includes a drive disk configured to be integral in rotation with a mast, an upper member at least partially located above the drive disk, a lower member at least partially located below the drive disk, and at least one link connecting the upper member and the lower member to the drive disk such that the drive disk drives the upper member and the lower member in rotation with the drive disk and that the upper member and the lower member are allowed to rotate differently with respect to the drive disk. The link having opposing end joints and a central joint that pivotally engages the drive disk. The system also having a gimbal device configured to drive a rotor hub and to allow gimbaling of the rotor hub with respect to a mast.

Abstract: A non-rotating universal joint for a helicopter drive unit has a spider defined by a ring, which has four connecting portions spaced 90° apart and engaging respective forks defining the ends of respective arms of two connecting members which, in use, are fixed with respect to a casing of an engine and a casing of a reduction gear; each fork and the corresponding connecting portion have respective through holes coaxial with one another and engaged by a screw; and dampers are interposed between the connecting portions of the ring and the shanks of the screws, and between the connecting portions of the ring and the forks.

Abstract: A method, system and a floating unit. The floating unit includes a propeller, a frame, a propeller motor that is configured to rotate the propeller about a first axis; wherein the propeller motor is coupled to the frame, a movable steering element; a controller, for controlling at least one of the propeller motor and the movable steering unit to affect at least one of a location and an orientation of the floating unit; and an interfacing module for coupling a payload to the floating unit and for receiving power from a connecting element that couples the floating unit to a ground unit; wherein the power received by the power interface is utilized to power the propeller motor and the controller.

Abstract: A multi-plate clutch is provided to couple a drive shaft from a gas turbine engine with a driven shaft of a lift fan. When disengaged the plates of the clutch can be physically separated. Plates of one shaft are coupled to the shaft with individual energy members that return the plates to a disengaged position. The plates of the other shaft can be coupled together with a travel member and/or lug key and energy devices between at least some of the plates.

Abstract: A model helicopter includes a drive mechanism comprising a battery powered motor comprising a toothed motor shaft; a main rotor comprising a main shaft aligned with the motor shaft and including a main gear; and a reduction gear assembly comprising a lower section including four first gears arranged around the motor shaft, and four second gears smaller than the first gears, each second gear fixed to top of the first gear; an upper section including four third gears and four fourth gears each being smaller than the third gear and fixed to top of the third gear, the fourth gears being meshed with the main gear; and four shafts each axially passing through the fourth gear, the third gear, the second gear, and the first gear so that a revolution of the motor shaft rotates the main shaft via the first gears, the fourth gears, and the main gear.

Abstract: A gas turbine engine includes a spool along an engine centerline axis which drives a gear train.

Abstract: A method for launching counter-measures from a launch opening of a dispenser mounted on an aeroplane. The counter-measures are launched in a direction obliquely forwards and downwards relative to the aeroplane. In order to facilitate launching of counter-measures, a low dynamic pressure is created permanently across the launch opening of the dispenser by a fixed element acting on the air stream. An arrangement for storing and launching counter-measures including an elongate body provided with at least one launch opening used for storing the counter-measures in compartments. The arrangement is designed to be mounted on an aeroplane with the longitudinal direction of the elongate body essentially coinciding with the flight direction of the aeroplane. The counter-measures are connected to a firing control unit for feeding firing signals to the counter-measures. The arrangement is designed to be mounted under the aeroplane.

Abstract: Electrically powered Vertical Takeoff and Landing (VTOL) aircraft are presented. Contemplated VTOL aircraft can include one or more electrical energy stores capable of delivering electrical power to one or more electric motors disposed within one or more rotor housings, where the motors can drive the rotors. The VTOL aircraft can also include one or more sustainer energy/power sources (e.g., batteries, engines, generators, fuel-cells, semi-cells, etc.) capable of driving the motors should the energy stores fail or deplete. Various VTOL configurations are presented including an all-battery purebred design, a light hybrid design, and a heavy hybrid design. The contemplated configurations address safety, noise, and outwash concerns to allow such designs to operate in built-up areas while retaining competitive performance relative to existing aircraft.

Abstract: The present invention relates to a power plant (10) having a single engine (13) together with both a main gearbox (MGB) suitable for driving the rotary wing (3) of a helicopter (1) and a tail gearbox (TGB) suitable for driving an anti-torque rotor (4) of a helicopter (1). The power plant (10) also includes a first electric motor (11) mechanically connected to said main gearbox (MGB) in order to be capable of driving said main gearbox (MGB), and a second electric motor (12) mechanically connected to said tail gearbox (TGB) in order to be capable of driving said tail gearbox (TGB).

Abstract: A gas turbine engine includes a fan section and a low spool that includes a low pressure compressor section. The low pressure compressor section includes eight (8) or less stages. A high spool includes a high pressure compressor section that has between eight to fifteen (8-15) stages. An overall compressor pressure ratio is provided by the combination of the low pressure compressor section and the high pressure compressor. A gear train is defined along an engine centerline axis. The low spool is operable to drive the fan section through the gear train.

Abstract: The invention relates to an electric drive (1) for an aircraft, especially for a helicopter (20) comprising at least one rotor (23) that is directly driven by a dynamoelectric machine (2). The dynamoelectric machine (2) is embodied in a duplex arrangement, a rotor (6) arranged in an air gap (12) comprising permanent magnets (13) on a carrier device (14), and the stators (4, 5) of the dynamoelectric machine (2) and/or the rotor (6) comprising coolant. A planetary gear (3) is especially provided between the driven rotor and the dynamoelectric machine (2), preferably in the axial extension of the dynamoelectric machine (2).

Abstract: Power takeoff means (46, 47, 60, 61) for this flight control system adjacent to the aircraft engines (41, 42) are all-electric and this makes it possible to reduce the length of hydraulic connections leading to actuators of the controlled devices (54, 55, 66). Electrical generators are of different types to minimize the risk of a generic failure. Characteristically of the invention, some generators (60, 61) have a constant voltage to frequency ratio and directly or almost directly control electro-hydraulic pumps (64, 65), without modifying the current characteristic and with no power electronics. Furthermore, system reconfigurations are made possible by additional electrical connections (69, 70, 71, 72) in case of failure.

Abstract: An aircraft propulsion unit includes a gas-turbine core engine 10 having at least one compressor, one combustion chamber and one turbine driving a main shaft 11. The main shaft 11 of the gas-turbine core engine 10 is operationally connected to at least two separate fans 6-9 via a mechanical drive connection, each of them being arranged beside the gas-turbine core engine 10.

Abstract: A hydraulic valve assembly includes a hydraulic valve, a control rod configured to be moved in translation along a first longitudinal axis by flight controls of an aircraft, a servocontrol, and an outlet rod configured to be moved in translation along a second longitudinal axis and configured to transfer a hydraulic fluid between the servocontrol and the hydraulic valve. A blocker device is configured to block a movement of the control rod when a difference between a first movement of the control rod and a second movement of the outlet rod exceeds a predetermined threshold and includes at least one moveable abutment configured to block the movement of the control rod when the predetermined threshold is reached and configured to move in longitudinal translation with the outlet rod, and a differential lever configured to mechanically link the control rod to the outlet rod and to cooperate with the at least one moveable abutment.

Abstract: A constant-velocity drive system for a rotary-wing aircraft rotor comprising a differential torque-splitting mechanism and a gimbal mechanism is disclosed. A rotary-wing aircraft having a rotary-wing aircraft rotor comprising a differential torque-splitting mechanism and a gimbal mechanism is disclosed.

Abstract: An aircraft that is housed within a gyroscope providing for improved flight stability that includes an inner hull which remains stationary within a rotating outer hull. A rotating plate includes a groove that traces a sinusoidal pattern around an inside periphery of the plate. Tracking arms have a portion that fit within the groove and another portion that connects to a piston. The rotation of the plate causes the tracking arms to travel within the sinusoidal groove and impart movement on the plurality of pistons which will intake air from above the upper surface of the inner hull configuration and create a negative pressure on the aircraft. The craft contains compression chambers which receive the air and which feed the various impeller thrusters which are rotational within three-fourths of a hemisphere.

Abstract: An exemplary lock out assembly includes a lock nut defined about an axis. The lock nut defines inner apertures along an inner diameter and outer apertures along an outer diameter. The inner apertures are at least partially threaded. The outer apertures are at least partially threaded. The inner apertures are radially spaced from the outer apertures.

Abstract: A lubrication system includes a reserve housing configured to retain a lubrication fluid. A supply line in fluid communication with the reserve housing is configured to provide pressurized lubrication fluid to the reserve housing. An overflow tube has an overflow port, the overflow tube being configured to prevent the volume of the lubrication fluid from exceeding a certain amount. A metering jet is configured to allow the lubrication fluid to flow from the reserve housing onto a component, such as a bearing, in the gearbox at a predetermined rate. The metering jet provides flow of the lubrication fluid onto the bearing even when the supply line no longer provides pressurized lubrication fluid to the reserve housing.

Abstract: A rotorcraft having multiple rotors, and wings that provide lift in forward flight, has a mechanical coupling between rotors that can be disengaged and optionally reengaged, during flight. The coupling, which can prevent a failure of one rotor from interfering with rotation of the other rotor(s), can be accomplished using many different types of devices, including for example, dog clutches, friction clutches, and collapsible clutches. Disengagement can range from being completely under control of an operator, to partially under operator control, to completely automatic. Among many other benefits, designing, manufacturing, fitting, retrofitting or in some other manner providing an aircraft with a device that can disengage rotation of one of the rotors from that of another one of the rotors during flight can be used to improve survivability in an emergency situation.

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 propeller-engine layout includes at least one engine (10) arranged in a tail (2) of a fuselage (1) of an aircraft and has at least one propeller arrangement, with a shaft of the propeller (6, 7) being connected to a shaft (21) of the engine (10) via at least one transfer shaft (12).

Abstract: An aircraft may have a fuselage having a longitudinal axis, 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, a first ducted fan rotatably mounted to the left wing, a second ducted fan rotatably mounted to the right wing, and an engine disposed in the fuselage which is connected to the first and second ducted fans.

Abstract: An aircraft (200) provided with a power plant (300) driving a rotary wing (60) and at least one propulsion means (7, 7?) possessing a propeller (6, 6?), said power plant (300) including at least one engine (1, 1?) having an outlet shaft (2, 2?) driving a drive train (400) in order to drive said rotary wing (60). The aircraft includes one coupling member (8, 8?) per engine (1, 1?) provided with a drive shaft (4, 4?) and with a first toothed wheel (13) that are coupled together by a clutch (15) and by a main controllable freewheel (25) arranged in parallel with the clutch, said drive shaft (4, 4?) driving at least one propulsion means (7, 7?) provided with brake means (9, 9?) for braking its propeller (6, 6?), said first toothed wheel (13) meshing with said drive train (400).

Abstract: A coaxial twin propeller model helicopter comprises a power assembly, a twin propeller rotation assembly, a tail rotation assembly, a motor assembly and a gear assembly. The twin propeller rotation assembly comprises a main rotation assembly and a minor rotation assembly. The power assembly comprises a motor assembly having a motor, a motor gear, and a fuel tank and a gear assembly having a main gear, a synchronous belt gear and a belt pulley. The main gear is engaged with the motor gear, and connected to the main rotation assembly. A lower part of the main gear is connected to the synchronous belt gear which connects to said minor rotation assembly. An upper end of the belt pulley is connected to said tail rotation assembly. A middle part of the belt pulley is engaged with the main gear. A lower end of the belt pulley is connected to the synchronous belt gear.

Abstract: A constant-velocity drive system for an aircraft rotor has a gimbal mechanism and a differential torque-combining mechanism. The gimbal mechanism has gimbals driven in rotation by a rotor mast about a mast axis, the gimbals providing for gimballing relative to the mast about gimbal axes generally perpendicular to the mast axis. The differential torque-combining mechanism is connected to the gimbal mechanism and configured to be driven in rotation about the mast axis by the gimbal mechanism. The differential mechanism is capable of gimballing relative to the mast about the gimbal axes, the differential torque-combining mechanism having an output component attached to a yoke of the rotor for driving the yoke in rotation with the differential torque-combining mechanism.

Abstract: An apparatus is provided and includes a rotatable shaft on which a first bevel gearbox is disposed, a layshaft on which a second bevel gearbox is disposed, the first and second bevel gearboxes being coupled to transmit shaft rotation to the layshaft, an accessory gearbox to which the layshaft and one or more accessories are coupled and a generator operably disposed on the layshaft between the second bevel gearbox and the accessory gearbox.

Abstract: Systems and methods for operating aircraft power systems are disclosed. A power system in accordance with one embodiment of the invention includes an aircraft auxiliary power unit and a controller coupled to the aircraft auxiliary power unit, with the controller being configured to automatically stop the auxiliary power unit while the auxiliary power unit is functioning normally. The controller can also be configured to automatically start the auxiliary power unit in-flight when power supplied or expected to be supplied to a subsystem of the aircraft has a non-zero value at or below a threshold value. The controller may also be configured to start the auxiliary power unit when a load or expected load on the subsystem meets or exceeds a threshold value.

Abstract: The invention relates to a power distribution system for an aircraft that comprises at least one engine, comprising at least one torque transmission means that is movably held and is connected to at least one shaft of the engine and extends from the shaft of the engine to an interior region of the aircraft fuselage, and at an end pointing away from the engine comprises at least one interface for introducing a torque. At the interface of the torque transmission means it is possible both to operate mechanical and to start the engine. Consequently in the aircraft it is possible to essentially do without converting mechanical energy to electrical energy and the subsequent use of electrical energy for generating mechanical power. This is particularly favourable in the case of twin-engine commercial aircraft that are used on short hauls.

Abstract: A vehicle including a fuselage having a longitudinal axis and a transverse axis, two Ducted Fan lift-producing propellers carried by the fuselage on each side of the transverse axis, a pilot's compartment formed in the fuselage between the lift-producing propellers and substantially aligned with one side of the fuselage, a payload bay formed in the fuselage between the lift-producing propellers and opposite the pilot's compartment, and two pusher fans located at the rear of the vehicle. Many variations are described enabling the vehicle to be used not only as a VTOL vehicle, but also as a multi-function utility vehicle for performing many diverse functions including hovercraft and ATV functions. Also described is an Unmanned version of the vehicle. Also described are unique features applicable in any single or multiple ducted fans and VTOL vehicles.

Abstract: A system for generating accessory power from a gas turbine engine is provided by the present invention. The system includes an electronic control device for monitoring at least one parameter which provides information about an incipient change in power demand, a control valve operated by the control device for supplying bleed air from the engine during a transient state in response to the at least one monitored parameter, and a pneumatically operated device for receiving the bleed air and for generating power to operate equipment onboard an aircraft. The pneumatically operated device may be an air turbine or a pneumatically integrated generator.

Abstract: A drive system for a high speed rotary-wing aircraft includes a combiner gearbox in meshing engagement with a main gearbox. The combiner gearbox is driven by one or more engines such that a main rotor system and a translational thrust system are driven thereby. The engine drives the combiner gearbox and thus the main gearbox through an overrunning clutch. The drive system permits the main rotor system RPM to be controlled by offloading power to the translational thrust system during a high speed flight profile.

Abstract: The present invention relates to a power plant (10) having a single engine (13) together with both a main gearbox (MGB) suitable for driving the rotary wing (3) of a helicopter (1) and a tail gearbox (TGB) suitable for driving an anti-torque rotor (4) of a helicopter (1). The power plant (10) also includes a first electric motor (11) mechanically connected to said main gearbox (MGB) in order to be capable of driving said main gearbox (MGB), and a second electric motor (12) mechanically connected to said tail gearbox (TGB) in order to be capable of driving said tail gearbox (TGB).

Abstract: The present invention relates to a propulsion device for an aircraft. The propulsion device comprises a first energy converter (4), a second energy converter (5) and a propulsion unit (1). The first energy converter (4) provides first propulsion energy, and the second energy converter (5) provides second propulsion energy. The first energy converter (4) and the second energy converter (5) are adapted to provide the first propulsion energy and the second propulsion energy to the propulsion device (1).

Abstract: The invention relates to a rotorcraft transmission system that comprises both a transmission shaft (21), and electromagnetic induction members (30, 31) of an active magnetic damper that extend around the shaft and that co-operate therewith to determine radial clearance (33). The system further comprises an additional radial damper (35) extending around the shaft with radial spacing (29) that is smaller than the radial clearance, such that in the event of the active magnetic damper failing, radial displacements of the shaft relative to its axis (22) are damped at least in part by the additional radial damper, so that damage to the induction members is limited or avoided.

Abstract: A multi-engine aircraft includes at least two first engines and a third engine located at a rear part of the fuselage, containing rear tail sections, along a vertical longitudinal plane of symmetry of the fuselage. The rear tail sections define a channel which is symmetrical with respect to the longitudinal plane of the fuselage. The third engine is arranged in the plane of symmetry of the channel corresponding to the longitudinal plane and is mounted on the upper part of the fuselage in a raised manner and in front of the tail sections, so that the outlet of the third engine is situated substantially at the inlet of the channel defined by the tail sections.

Abstract: A light weight, split torque, geared power transmission having a reduced number of gears and useful for, e.g., transmitting power from an engine to the main rotor of a rotorcraft, includes a ring gear having a pair of oppositely facing angular bevel gears respectively formed on opposite sides of a medial plane thereof, and a first pinion having an elongated drive shaft extending through the medial plane of the ring gear and disposed at an oblique angle relative thereto. The first pinion includes a pair of cylindrical gears mounted coaxially on the drive shaft and respectively disposed in conjugate meshing engagement with respective ones of the bevel gears of the ring gear. An elongated annular output shaft made of a strong, light weight composite material is coupled to a circumferential periphery of the ring gear.

Abstract: Systems and methods for operating aircraft power systems are disclosed. A power system in accordance with one embodiment of the invention includes an aircraft auxiliary power unit and a controller coupled to the aircraft auxiliary power unit, with the controller being configured to automatically stop the auxiliary power unit while the auxiliary power unit is functioning normally. The controller can also be configured to automatically start the auxiliary power unit in-flight when power supplied or expected to be supplied to a subsystem of the aircraft has a non-zero value at or below a threshold value. The controller may also be configured to start the auxiliary power unit when a load or expected load on the subsystem meets or exceeds a threshold value.

Abstract: A propeller or propeller drive, the propeller being carried by a support shaft (3) which projects from the body (53) of the missile or vehicle and about which the propeller rotates, the drive unit for the propeller being situated in front of the propeller (23) in the propulsion direction or in the propeller spinner (1). According to the invention it is provided that the drive unit has at least one gas turbine (55), optionally a multiple-stage gas turbine, which is secured to the drive shaft and which drives it or jointly rotates with it, and that the drive shaft (6) is traversed by the support shaft (3) or is mounted thereon in a rotating manner and drives the propeller (23), optionally by way of a gear mechanism (14) situated between the turbine (55) and the propeller (23).

Abstract: A fixed wing Vertical Take-Off and Landing (VTOL) aircraft for use as a Personal Air Vehicle (PAV) or unmanned vehicle. A first double-ended drive shaft engine is mounted sideways in the front of the fuselage to serve a first pair of ducted fans mounted at the ends of the front wing. A second double-ended drive shaft engine is mounted sideways in the rear of the fuselage to serve a second pair of ducted fans mounted on the rear fuselage. The ducted fans are rotatable from a horizontal orientation to a vertical orientation to permit the aircraft to take off and land as a VTOL or conventional aircraft, and to be flown as a conventional aircraft. A parachute is provided with inflation assistance to permit rapid low altitude deployment for a controlled descent of the aircraft in an emergency.