264results about "Depending on number of power plants" patented technology

An unmanned air vehicle is provided, which includes an airframe and a parallel hybrid-electric propulsion system mounted on the airframe. The parallel hybrid-electric propulsion system includes an internal combustion engine and an electric motor. A hybrid controller is configured to control both the internal combustion engine and the electric motor. A propeller is connected to a mechanical link. The mechanical link couples the internal combustion engine and the electric motor to the propeller to drive the propeller. An alternate unmanned air vehicle includes a second propeller driven by the electric motor. In this alternate unmanned air vehicle, the internal combustion engine is decoupled from the electric motor.

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.

An unmanned aircraft includes a propulsion system having a diesel or kerosene internal combustion engine and a charger device for engine charging. The propulsion system can be a hybrid propulsion system or a parallel hybrid propulsion system.

A vertical take-off and landing aircraft including a wing structure including a wing, a rotor operatively supported by the wing, and a hybrid power system configured to drive the rotor, the hybrid power system including a first power system and a second power system, wherein a first energy source for the first power system is different than a second energy source for the second power system.

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.

A multi-engine system 10 for use on an aircraft 14 is disclosed. The system 10 may comprise a first aircraft engine 16A and a second aircraft engine 16B configured to drive at least one device 12 of the aircraft 14. A starter 20 may be coupled to one of the first and second aircraft engines 16A, 16B to assist starting of the one of the first and second aircraft engines 16A, 16B. An energy source 24 may be configured to deliver energy to the starter 20 at a rapid rate during rapid starting of the one of the first and the second aircraft engines 16A, 16B, the rapid rate being higher than a regular rate used during regular starting of the one of the first and the second aircraft engines 16A, 16B. The energy source 24 may comprise at least one supercapacitor.

A method of managing an engine failure on a rotary wing aircraft (1) having a hybrid power plant (5) with at least two fuel-burning engines (13, 13′), at least one electric machine (12), and a main gearbox (11). Said aircraft (1) also has electrical energy storage means (14) and a main rotor (2) mechanically connected to said hybrid power plant (5). In said method, during each flight, the operation of said engines (13, 13′) is monitored in order to detect a failure of any one of them, and then once a failure of one of said engines (13, 13′) has been detected, said electric machine (12) is controlled, if necessary, to deliver auxiliary power We in order to avoid a deficit appearing in the total power WT of said hybrid power plant (5), thereby enabling the pilot of said aircraft (1) to fly said aircraft (1) safely without degrading said hybrid power plant (5).

A vertical take-off aircraft with a propulsion drive for generating a driving force being effective in a horizontal direction and with a lift drive for generating a lifting force being effective in a vertical direction includes a motor for providing mechanical energy for the propulsion drive and a first generator for providing electrical energy for the lift drive. Moreover, the aircraft includes an exhaust gas turbocharger for the motor with a first turbine being driven by an exhaust gas flow of the motor, wherein the first turbine is configured to provide mechanical energy for the propulsion drive.

The system of the present application includes an engine and pylon arrangement for a tilt rotor aircraft in which the engine is fixed in relation to a wing portion of the aircraft, while the pylon is rotatable. The pylon supports a rotor hub having a plurality of rotor blades. Rotation of the pylon allows the aircraft to selectively fly in a helicopter mode and an airplane mode, as well as any combination thereof.

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.

The system of the present application includes an engine and pylon arrangement for a tilt rotor aircraft in which the engine is fixed in relation to a wing portion of the aircraft, while the pylon is rotatable. The pylon supports a rotor hub having a plurality of rotor blades. Rotation of the pylon allows the aircraft to selectively fly in a helicopter mode and an airplane mode, as well as any combination thereof.

Provided are alternative hybrid transmission systems for marine, or two wheeled land vehicles, as well as propulsion systems and vehicles comprising such transmission systems, to improve various propulsion systems using a combination of at least two power sources with the option for simultaneous or alternating power input from two or more power sources, while providing desired characteristics or components. Such characteristics or components can include, but are not limited to: power, torque, acceleration, cruising speed or power, fuel efficiency, battery charging, endurance, power sizing, weight, capacity, efficiency, speed, mechanically and / or electrically added system requirements, design, fuel selection, functional design, structural design, lift to drag ratio, weight, and / or other desired characteristic or component.

Rotary-wing aircraft or fixed-wing aircraft consisting of one or a plurality of rotors and / or one or a plurality of airscrews being caused to rotate by means of at least one shaft, the said aircraft comprising a distributed electric motor unit configured to ensure the propulsion and / or the lifting of the said aircraft by causing the said shaft to rotate, the motor unit being a distributed electric motor unit connected directly to the rotating shaft, no mechanism for the transmission of movement being interposed between the said CND unit and the said shaft, the distributed electric motor unit being composed of a plurality of stacked electric motor elements, each said electric motor element being connected directly to the rotating shaft and consisting of at least one fixed stator and at least one moving rotor capable of connection to the said rotating shaft in order to transmit to it a mechanical power of the following kind where: Pr=the nominal mechanical power necessary for the propulsion and / or for the lifting of the said aircraft, Pim=the maximum mechanical power capable of being delivered by the electric motor element on level i, to the rotating shaft, for instance Pim<Pr, Ki=the power derating of the electric motor element on level i, for instance 0≦ki≦1, Ki being a variable that is adjustable as a function of the undamaged number of electric motor elements and / or as a function of the nominal mechanical power necessary for the propulsion and / or for the lifting of the said aircraft at a moment in time t, n=the number of electric motor elements comprising the distributed electric, motor unit, for instance n≧2,—the axes of rotation of the moving rotors and of the rotating shaft being coaxial.

A method of assisting a pilot of a single-engined rotary wing aircraft (1) during a stage of flight in autorotation, said aircraft (1) including a hybrid power plant having an engine (13), an electric machine (12), and a main gearbox (11). Said aircraft (1) also includes electrical energy storage means (14) and a main rotor (2) mechanically connected to said hybrid power plant (5). According to said method, while in flight, the operation of said engine (13) is monitored in order to detect a failure thereof, in particular by monitoring for a drop of power on said main rotor (2), and then in the event of a failure of said engine (13) being detected, said electric machine (12) is controlled to deliver auxiliary power We to said main rotor (2), thereby enabling the pilot of said aircraft (1) to be assisted in maneuvering said aircraft (1) during said stage of flight in autorotation following said failure.

An extremely large aircraft which is suitable for overseas cargo transport and which includes a fuselage defining a central storage cavity, a wing assembly defining a pair of wing storage cavities, an altitude control system, and a plurality of independently steerable landing gear units. The central storage cavity has a length, height and width of at least 100 feet, at least 16 feet and at least 24 feet, respectively. The wing assembly has a wingspan of at least 300 feet and is configured with a moderate aspect ratio to permit both ground-effect and high altitude operation. The altitude control system controls the aircraft in ground effect such that the aircraft is maintained at about a predetermined altitude. The landing gear units are coupled to the fuselage and are arranged in at least two discrete columns and at least ten discrete rows. The central storage cavity and the wing storage cavities are configured to receive cargo including intermodal re-usable cargo containers.

A method of managing an engine failure on a rotary wing aircraft (1) having a hybrid power plant (5) with at least two fuel-burning engines (13, 13′), at least one electric machine (12), and a main gearbox (11). Said aircraft (1) also has electrical energy storage means (14) and a main rotor (2) mechanically connected to said hybrid power plant (5). In said method, during each flight, the operation of said engines (13, 13′) is monitored in order to detect a failure of any one of them, and then once a failure of one of said engines (13, 13′) has been detected, said electric machine (12) is controlled, if necessary, to deliver auxiliary power We in order to avoid a deficit appearing in the total power WT of said hybrid power plant (5), thereby enabling the pilot of said aircraft (1) to fly said aircraft (1) safely without degrading said hybrid power plant (5).

An aircraft (1) having a rotary wing (2) and at least one main gearbox (5) for driving rotation of said rotary wing (2), said aircraft (1) having a first main engine (11) and a second main engine (12) driving said main gearbox (5), said aircraft (1) being provided with a main regulation system (15) regulating the first main engine (11) and the second main engine (12) in application of a setpoint that is variable. A secondary engine (21) is also capable of driving said main gearbox (5), said aircraft (1) having a secondary regulation system (25) that regulates the secondary engine (21) in application of a setpoint that is constant and that is independent of said main regulation system (15).

Systems and methods perform power conversion and distribution. A system for power conversion and distribution, according to one embodiment comprises: a first generator substation (70) receiving power from a first electrical generator (10); a first main substation (100) for distributing converted power to at least one load, said load being local to the first main substation (100); a ring bus including ring bus feeders (1000, 2000); and at least one satellite substation (700) for powering a remote load, wherein the ring bus feeders (1000, 2000) connect the first generator substation (70) and the first main substation (100) in a ring arrangement, and the at least one satellite substation (700) receives power from the ring bus through a branch substation (500).

A hybrid power drive system for an aircraft comprises a rotor that receives power and a first power drive sub-system including at least one engine in connection with the rotor is configured to provide a first power to the rotor. Further, the hybrid power drive system also includes a second power drive sub-system connected in parallel to the first power drive sub-system. The second power drive sub-system is configured to provide a second power to the rotor a second power drive sub-system connected in parallel to the first power drive sub-system and configured to provide a second power to the rotor when the first power provided by the first power drive sub-system is less than a power demand of the rotor.

A motorized aircraft is controlled by a drive control means (120) that rotationally drives a propulsion system propeller. The propulsion system propeller is driven by a plurality of electric motors (130). Total output of the plurality of electric motors (130) and required output satisfy the following relationship: Pmax(n−1) / n>Preq (where Pmax is the total output (kW) of the electric motors, n is the number of electric motors, and Preq is the required output (kW)).

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.

A propulsion system includes an engine disposed within a fuselage, a first gearbox coupled to the engine, a wing member having an upper wing skin and torque box formed by a first rib, a second rib, a first spar and second spar, a drive shaft coupled to the first gearbox and disposed within the wing member, a second gear box coupled to the drive shaft and disposed outboard from the second rib or inboard from the first rib, and a rotatable proprotor coupled to the second gear box. The rotatable proprotor includes a plurality of rotor blades, a rotor mast having a mast axis of rotation, and a proprotor gearbox coupled to the rotor mast. The proprotor gearbox is rotatable about a conversion axis, which intersects with the mast axis of rotation at a point within a central region of the torque box and above the upper wing skin.

The invention discloses a driving mechanism of a hybrid power helicopter. The driving mechanism of the hybrid power helicopter comprises a fuel tank, an engine, a clutch, a first locker, a planetary gear mechanism, an energy storage device, a generator / electric motor, a second locker, a second generator / electric motor, a torque coupler, a tail end gear box, a tail rotor subsystem, an swashplate and a rotor wing subsystem, wherein the clutch and the first locker are mounted between the engine and the planetary gear mechanism, and the second locker is mounted between the generator / electric motor and the planetary gear mechanism. Two power sources comprising a basic power source and an auxiliary power source are utilized, the two power sources are reasonably matched and selected when driving the helicopter, accordingly, the problems that the power source of a conventional helicopter is single, a drive system is complicated, a drive device is heavy, mechanical failure easily happens and the like are solved, and the safety and the maneuverability of the helicopter are improved.

A hybrid-electric propulsion system for an aircraft includes a propulsor and a turbomachine. The turbomachine includes a high pressure turbine drivingly coupled to a high pressure compressor through a high pressure spool. The hybrid electric propulsion system further includes an electrical system including a first electric machine, a second electric machine, and an electric energy storage unit electrically connectable to the first and second electric machines. The first electric machine is coupled to the high pressure spool of the turbomachine and the second electric machine is coupled to the propulsor for driving the propulsor to provide a propulsive benefit for the aircraft. The hybrid electric propulsion system also includes a controller configured to provide electrical power from an electric power source to the first electric machine to drive the first electric machine to start, or assist with starting, the turbomachine.

The invention relates to a rotary- or fixed-wing aircraft comprising one or more rotors and / or one or more propellers, in which the rotor(s) and / or propeller(s) are rotated at a variable or constant speed by at least one shaft (Rp1, Rp2, Rp, RAC, H), and the aircraft includes a power unit (GEMD) configured to propel and / or lift the aircraft by rotating the shaft. The aircraft is characterised in that: the power unit (GEMD) is a distributed electric power unit formed by multiple stacked electric power elements (Ee1, Ee2, Een), each element being adapted to confer, on the rotation shaft (Rp1, Rp2, Rp, RAC, H), a fraction of the total power necessary in order to propel and / or lift the aircraft; the distributed electric power unit (GEMD) is in direct engagement with the rotation shaft (Rp1, Rp2, Rp, RAC, H), with no movement transmission mechanism being inserted between the unit and the shaft; each electric power element (Ee1, Ee2, Een) is directly connected to the rotation shaft (Rp1, Rp2, Rp, RAC, H) and comprises at least one fixed stator (St), at least one moving rotor (Rt) and at least one mechanical or electromagnetic free wheel (RI) in direct engagement with the rotation shaft, said moving rotor (Rt) co-operating with the free wheel (RI) such as to be coupled to the rotation shaft during normal operation of the electric power unit (Ee1, Ee2, Een) and to be disconnected from the shaft in the event of the failure of the electric power element; and the rotation shafts of the moving rotors (Rt), the free wheels (RI) and the rotation shaft (Rp1, Rp2, Rp, RAC, H) are coaxial.

A propulsion system includes a propulsor having a driveshaft, an electric machine coupled to the driveshaft of the propulsor, and a combustion engine having an output shaft. The propulsion system additionally includes a one-way clutch operable with at least one of the driveshaft of the propulsor and the output shaft of the combustion engine. The one-way clutch allows for a differential angular velocity of the driveshaft relative to the output shaft in a first circumferential direction and prevents a differential angular velocity of the driveshaft relative to the output shaft in a second circumferential direction.

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).

The invention relates to an architecture of a propulsion system of a multiple-engine helicopter comprising turboshaft engines (5, 6), characterised in that it comprises: at least one hybrid turboshaft engine (5) that is capable of operating in at least one standby mode during a stable flight of the helicopter, the other turboshaft engines (6) operating alone during this stable flight; an air turbine (30) that is mechanically connected to the gas generator (17) of the hybrid turboshaft engine (5) and is suitable for rotating said gas generator (17); means for withdrawing pressurised air from the gas generator (27) of a running turboshaft engine (6); and a duct (31) for routing this withdrawn air to said air turbine (30).

An electrical architecture for an aircraft having a main electrical machine connected to a fuel burning engine and a secondary electrical machine connected to a power transmission assembly. A high-voltage electrical master box is connected by a first line and by a second line to a multifunction converter, the high-voltage electrical master box connecting the first line to the secondary electrical machine, and the second line to at least the main electrical machine and the secondary electrical machine. The multifunction converter includes a supervisor connected to an avionics system and to a control system controlling the engine, and to a controller of the high-voltage electrical master box.