58results about "Depending on power plant type" patented technology

Methods, apparatus, systems, and articles of manufacture are disclosed for a hydrogen fuel distribution system (100) with integrated thermal management. An example fuel distribution system (100) with integrated thermal management comprises a hydrogen fuel distributor (402) including a first tank (406) and a second tank (408). The hydrogen fuel distributor (402) circulates hydrogen fuel to a first fuel cell (606) and a second fuel cell (608). A coolant distributor (702) includes a heat exchanger (416). The coolant distributor (702) regulates a temperature of the fuel distribution system (100). A compressed air distributor (706) provides compressed air from a propulsor (730) to the first fuel cell (606) and the second fuel cell (608). An electrical distributor (602) transports electricity from the first fuel cell (606) and the second fuel cell (608) to an electric motor (726). The electric motor (726) drives the propulsor (730).

A method for controlling a thermal and electrical power plant for setting in motion at least one rotary member of a rotorcraft, the power plant comprising at least one heat engine and an electrical system provided with at least one electric machine. The method comprises: selecting, with a selector, an operating mode chosen from several operating modes; determining a density altitude and comparing, with a controller, the current density altitude and a threshold density altitude; and controlling, with the controller, the at least one electric machine depending on at least the chosen operating mode as well as the comparison and a necessary power to be supplied to the power transmission system.

A propulsion system for an aircraft includes a first open propulsor rotor, a second open propulsor rotor, a turbine engine and a drive system. The second open propulsor rotor is next to and downstream of the first open propulsor rotor. The turbine engine includes a flowpath, a compressor section, a combustor section, a turbine section and a rotating assembly. The rotating assembly includes a turbine rotor in the turbine section. The rotating assembly is configured to drive rotation of the first open propulsor rotor during a first mode and a second mode. The drive system is discrete from the rotating assembly. The drive system is configured to drive rotation of the second open propulsor rotor during the first mode. The second open propulsor rotor is rotationally fixed during the second mode.

Modular engine (1) for missiles, comprising: an engine inlet (10); a gas turbine (12); a payload turbine (14); at least one propulsor (20) and a nozzle assembly (16), wherein an outlet of the engine inlet (10) is connected to an inlet of the gas turbine (12), and wherein an outlet of the gas turbine (12) is selectively connectable to an inlet of the nozzle assembly (16) and / or to an inlet of the payload turbine (14) for activating and deactivating the payload turbine (14) for driving the at least one propulsor (20), characterized in that the engine (1) is designed to deactivate the payload turbine by jettisoning the payload turbine (14) and / or the at least one propulsor (20).

A lightweight multi-condition hybrid parallel shaft transmission for aviation belongs to the technical field of lightweight multi-mode hybrid parallel shaft transmissions for aviation. Its features include an output shaft and an input shaft. The output shaft side surface is fitted with an integrated propeller, oil seal, C2 clutch, and driven gear from left to right. A controller is connected to the center of the right end of the output shaft. A C1 clutch is connected to the left end of the input shaft. An engine is fixedly connected to the center of the right end of the input shaft. A motor shaft is connected to the left end of the C1 clutch. The motor shaft is fitted with a motor, drive motor, resolver, drive gear, and oil pump gear from right to left. Several bearings are inserted into the side surfaces of the output shaft, input shaft, and motor shaft, facilitating disassembly, assembly, maintenance, and repair, reducing the cost per use throughout the product's lifespan, further increasing product redundancy, and reducing the failure rate.

In a first aspect, described herein is a direct drive electric aircraft propulsion wherein the propulsion rotor torque is decoupled from the primary proprotor mast moment forces. A hub shaft locates the propulsion proprotor assembly in space relative to the aircraft nacelle, while a motor torque coupler transfers torque from the electric motor to the propulsion rotor while resolving a negligible amount of mast moment through the electric motor.

An assembly for an aircraft includes a first propulsion system (24B) and a second propulsion system (24A). A first open propulsor rotor (60) is configured to rotate in a first rotational direction about a first axis (46). A first engine core (102) includes a first rotating structure (116) with a first turbine rotor (107). The first rotating structure (116) is configured to rotate in the first rotational direction about the first axis (46) and drive rotation of the first open propulsor rotor (60) through a first geartrain (58). A second open propulsor rotor is configured to rotate in a second rotational direction about a second axis that is opposite the first rotational direction. A second engine core includes a second rotating structure with a second turbine rotor. A second rotating structure is configured to rotate in the first rotational direction about the second axis and drive rotation of the second open propulsor rotor through a second geartrain.

The invention relates to an aircraft comprising: - a fuselage (32), - a wing (34), - at least one hydrogen-powered turbomachine (40) generating thrust at a propulsion unit (36) located away from the fuselage (32), - at least one fuel tank (46) positioned in the fuselage (32) and configured to store hydrogen in a cryogenic state, - at least one hydrogen supply system connecting the fuel tank (46) and the turbomachine (40) comprising: o at least one pump (50) positioned in the fuselage (32) near the fuel tank (46), o at least one hydrogen heating system (52) positioned in the fuselage (32) near the pump (50). This solution makes it possible to reduce the length of the complex double-walled conduits configured to channel the hydrogen in a cryogenic state between the fuel tank and the hydrogen heating system.

An aircraft assembly includes a fan case, a core case, an accessory gearbox, one or more electric machines, a first electric device and an electrical system. The core case is radially inboard of the fan case. The accessory gearbox is mounted with the fan case or the core case. The one or more electric machines are mounted to the accessory gearbox. The accessory gearbox is dedicated to the one or more electric machines. The first electric device is mounted with the fan case or the core case. The electrical system is electrically coupled to the one or more electric machines and the first electric device. The electrical system is configured to receive a first current of electricity from at least one of the one or more electric machines. The electrical system is configured to provide a second current of electricity to the first electric device.

An assembly for an aircraft includes a first propulsion system and a second propulsion system. A first open propulsor rotor is configured to rotate in a first rotational direction about a first axis. A first engine core includes a first rotating structure with a first turbine rotor. A first rotating structure is configured to rotate in the first rotational direction about the first axis and drive rotation of the first open propulsor rotor through a first geartrain. A second open propulsor rotor is configured to rotate in a second rotational direction about a second axis that is opposite the first rotational direction. A second engine core includes a second rotating structure with a second turbine rotor. A second rotating structure is configured to rotate in the first rotational direction about the second axis and drive rotation of the second open propulsor rotor through a second geartrain.

A system for an aircraft is provided, the aircraft including a tiltable prop-rotor tiltable between a lift position for providing lift to the aircraft and a forward flight position for providing forward thrust to the aircraft, the system including at least one actuator for adjusting a tilt angle of the tiltable prop-rotor, and at least one passive damper connected to the tiltable prop-rotor and configured to limit a rate of change of the tilt angle of the tiltable prop-rotor.

A decoupling transmission system (2) for an aircraft propulsion unit (1) comprises: a first rotating element (3) carrying a ramp (131), a second rotating element (7), and coupling means (116) for these; an actuator (40) comprising a plunger (140) movable from a rest position to an operating position to actuate the ramp (131) towards a decoupling position of the rotating elements (3, 7), and a cam (50). The plunger (140) carries a cam follower (56) held in contact with a surface (54) of the cam by means of a plunger return means (144). A displacement of the cam (50) by a first angular distance (θ1) in a first direction (D1) moves the plunger (140) from its rest position to its operating position.A displacement of the cam (50) by a second smaller angular distance (θ2), and in a second opposite angular direction (D2), moves the plunger (140) from its rest position to its operating position. See Figure 3 for abbreviations.

An open propulsor rotor is configured to rotate about a first axis. An open guide vane structure includes a plurality of open guide vanes arranged circumferentially about a second axis offset from the first axis. The open guide vanes include a first open guide vane. A turbine engine includes a flowpath, a compressor section, a combustor section, a turbine section and a rotating assembly. The flowpath extends through the compressor section, the combustor section and the turbine section from an airflow inlet into the flowpath to a combustion products exhaust from the flowpath. The airflow inlet is axially aligned with or located forward of the first open guide vane along the second axis. The rotating assembly includes a turbine rotor disposed in the turbine section. The rotating assembly is configured to drive rotation of the open propulsor rotor through an offset geartrain.

A propulsion system for an aircraft includes an open propulsor rotor, a turbine engine and a thermal management system. The turbine engine includes an engine core and an engine flowpath. The engine core includes a compressor section, a combustor section and a turbine section. The engine flowpath extends longitudinally though the compressor section, the combustor section and the turbine section from an airflow inlet into the engine flowpath to a combustion products exhaust from the engine flowpath. The thermal management system is configured to manage heat energy generated by the propulsion system during operation of the turbine engine. The thermal management system includes an electric boost compressor, a heat exchanger and a system flowpath disposed outside of the engine core. The system flowpath extends longitudinally through the electric boost compressor and the heat exchanger from an airflow inlet into the system flowpath to an airflow exhaust from the system flowpath.

A propulsion system (20) for an aircraft includes an open propulsor rotor (38), a turbine engine (32) and a thermal management system (36). The turbine engine (32) includes an engine core (82) and an engine flowpath (76). The engine core (82) includes a compressor section (71), a combustor section (72) and a turbine section (73). The engine flowpath (76) extends longitudinally though the compressor section (71), the combustor section (72) and the turbine section (73) from an airflow inlet (78) into the engine flowpath (76) to a combustion products exhaust (80) from the engine flowpath (76). The thermal management system (36) is configured to manage heat energy generated by the propulsion system (20) during operation of the turbine engine (32). The thermal management system (36) includes an electric boost compressor (114), a heat exchanger (116) and a system flowpath (110) disposed outside of the engine core (82). The system flowpath (110) extends longitudinally through the electric boost compressor (114) and the heat exchanger (116) from an airflow inlet (122) into the system flowpath (110) to an airflow exhaust (124) from the system flowpath (110).

Systems, apparatuses, and methods for overcoming the disadvantages of current air transportation systems that might be used for regional travel by providing a more cost effective and convenient regional air transport system. In some embodiments, the inventive air transport system, operational methods, and associated aircraft include a highly efficient plug-in series hybrid-electric powertrain (specifically optimized for aircraft operating in regional ranges), a forward compatible, range-optimized aircraft design, enabling an earlier impact of electric-based air travel services as the overall transportation system and associated technologies are developed, and platforms for the semi-automated optimization and control of the powertrain, and for the semi-automated optimization of determining the flight path for a regional distance hybrid-electric aircraft flight.

The invention relates to a hybrid drive system (1) for an aircraft. A gas turbine engine (2) is designed to drive an engine output shaft (10) which can be connected to a planetary gearbox (4) via a freewheel (16). An electric machine (3) is designed to drive a machine output shaft (18) which can be operatively connected to the planetary gearbox (4) via a clutch (20). A gear unit (5) comprises an input shaft (11) and an output shaft (14), the input shaft (11) extending radially offset with respect to the output shaft (14) and coaxially with respect to the engine output shaft (10). The input shaft (11) can be connected to the engine output shaft (10) via the freewheel (16), and the output shaft (14), which is arranged coaxially with respect to the machine output shaft (18), can be connected to the planetary gearbox (4). Alternatively, the output shaft (14) may be connectable, via the freewheel (16), to the planetary gearbox (4) by means of a further shaft (19) which extends coaxially with respect to the machine output shaft (18), and the input shaft (11) may be connected coaxially to the engine output shaft (10).

To provide a power distribution circuit for aircraft and an electric aircraft. [Solution] The power distribution circuit for an aircraft comprises a plurality of batteries, a plurality of inverters, each configured to be coupled to each of the plurality of batteries and to generate a plurality of input phases, a plurality of interphase transformers, each receiving one or more of the plurality of input phases and generating a combined single drive phase, and a motor configured to receive the combined single drive phase from each of the plurality of interphase transformers.

A propulsion system (50) is disclosed. The propulsion system (50) includes a first propeller (52), a second propeller (54), and a third propeller (56). The first propeller (52), the second propeller (54), and the third propeller (56) are arranged to rotate about a common axis and the second propeller (54) is disposed between the first propeller (52) and the third propeller (56). The first and third propellers (52, 56) are configured to rotate about the common axis in a first direction (A) and the second propeller (54) is configured to rotate about the common axis in a second direction (B) opposite to the first direction (A). A first motor (60) may be coupled to the first and third propellers (52, 56) and a second motor (64) may be coupled to the second propeller (54). A first shaft (58) and second shaft (62) may be arranged along the common axis, wherein the first and third propellers (52, 56) are coupled to the first shaft (58) and the second propeller (54) is coupled to the second shaft (62).

The invention discloses a long-endurance hydraulic power multi-rotor unmanned aerial vehicle, and relates to the field of long-endurance unmanned aerial vehicles, the long-endurance hydraulic power multi-rotor unmanned aerial vehicle comprises a vehicle body, a front arm rod is fixedly connected to the front end of the vehicle body, side arm rods are fixedly connected to the two sides of the vehicle body, a rear arm rod is fixedly connected to the rear end of the vehicle body, an inner cavity is formed in the rear arm rod, and a ventilation hole is formed in the top surface of the inner cavity; the unmanned aerial vehicle has the advantages that the unmanned aerial vehicle can be quickly braked, can be quickly stabilized after arriving at a target point, and can also be quickly decelerated in an emergency, so that the unmanned aerial vehicle can be suddenly dropped to a hovering state from a high-speed state within an extremely short distance and time, and the working efficiency and the flexibility of route planning are greatly improved; meanwhile, the anti-mistaken touch effect is achieved, the unmanned aerial vehicle is prevented from falling to the ground due to height loss, the control risk is reduced, the safety of the unmanned aerial vehicle is improved, the parallel pipelines can select circulation paths according to requirements, complex mechanisms do not need to participate in adjustment, and the control and material cost is reduced.

The present application relates to systems and methods for implementing a regional air transportation network using hybrid-electric aircraft. Systems, devices and methods for overcoming the shortcomings of current air transportation systems available for regional travel by providing a more cost-effective and convenient regional air transportation system. In some embodiments, the present air transportation system, method of operation and related aircraft include: a highly efficient plug-in series hybrid electric powertrain optimized specifically for aircraft operating within regional ranges; a forward-compatible range-optimized aircraft design based on electric air travel services that can be influenced earlier as the overall transportation system and related technologies evolve; a platform for semi-automated optimization and control of the powertrain and determination of the flight path for semi-automated optimization of regional distance hybrid-electric aircraft flights.

An aircraft assembly includes an engine core, a first electrical power system and a second electrical power system. The first electrical power system includes a first electric machine and a first electric machine controller. The first electric machine is operatively coupled to the engine core. The first electric machine controller is electrically coupled to the first electric machine. The first electric machine controller is configured to control operation of the first electric machine. The second electrical power system is electrically discrete from the first electrical power system. The second electrical power system includes a second electric machine and a second electric machine controller. The second electric machine is operatively coupled to the engine core. The second electric machine controller is electrically coupled to the second electric machine. The second electric machine controller is configured to control operation of the second electric machine.

A hybrid propulsion system (1) for an aircraft is described. A gas turbine engine (2) is configured to drive an engine output shaft (10), which can be connected to a planetary gearbox (4) via a freewheel (16). An electric machine (3) is configured to drive a machine output shaft (18), which can be operatively connected to the planetary gearbox (4) via a coupling (20). A gearbox unit (5) comprises an input shaft (11) and an output shaft (14), wherein the input shaft (11) is radially offset from the output shaft (14) and coaxial with the engine output shaft (10). The input shaft (11) can be connected to the engine output shaft (10) via the freewheel (16), and the output shaft (14), which is arranged coaxially with the machine output shaft (18), can be connected to the planetary gearbox (4).Alternatively, the output shaft (14) can be connected to the planetary gear (4) via the freewheel (16) with another shaft (19) which runs coaxially to the machine output shaft (18), and the input shaft (11) can be connected coaxially to the engine output shaft (10).

An aircraft propulsion system (1) comprising: a propulsion unit (2) mounted on a propulsion shaft (3); a gas turbomachine (4) comprising a compressor (41) and a turbine (43) mounted on a drive shaft (5); a transmission system (6) configured to define a transmission ratio (K) between a propulsion rotational speed (V3) of the propulsion shaft (3) and a motor rotational speed (V5) of the drive shaft (5); a first electric machine (M1) configured to supply or draw control power from the transmission system (6) so as to modify the transmission ratio (K); and a second electric machine (M2) connected to the propulsion shaft (3) or the drive shaft (5) and configured to supply or draw compensating power from the propulsion shaft (3) so as to compensate for residual power induced by the application of the control power. Abstract figure: Figure 2

A tractor aircraft power plant for propelling an aircraft is described. The tractor aircraft power plant is configured to improve serviceability of swirl-recovery vanes. The tractor aircraft power plant includes an intermittent combustion engine such as a Wankel engine operable to generate motive power. A bladed rotor is drivingly coupled to the Wankel engine and is disposed forward of the Wankel engine relative to a propulsion direction. A gearbox drivingly couples the bladed rotor to the Wankel engine. A plurality of vanes are angularly distributed about a rotation axis of the bladed rotor and are non-rotatable about the rotation axis of the bladed rotor. The vanes are disposed aft of the bladed rotor to interact with a flow of air propelled by the bladed rotor during rotation of the bladed rotor. The vanes are mounted to the gearbox to facilitate serviceability of the vanes.

To provide an electric aircraft having a DC bus and a plurality of batteries each coupled in parallel to the DC bus.SOLUTION: At least one switch is coupled in series between at least one battery of the plurality of batteries and a DC bus, and a plurality of inverter circuits are each coupled in parallel to the DC bus. A plurality of motors are coupled to the plurality of inverter circuits, respectively. In various embodiments, an electric aircraft further includes a plurality of switches, each switch coupled in series between respective batteries of the plurality of batteries.SELECTED DRAWING: Figure 1

A propulsion system for an aircraft includes a gas turbine engine, an electrical assembly, a powertrain, and a controller. The gas turbine engine includes a rotational structure and a second rotational structure each including a turbine rotor. The electrical assembly includes an electric machine and a control unit. The control unit is electrically connected to the electric machine. The powertrain includes a differential geartrain configured to couple the rotational structure and the second rotational structure together to the electric machine. The controller is connected in communication with the control unit. The controller is configured to maintain a predetermined power transfer ratio between a spool mechanical power transfer of the second rotational structure and an electric machine mechanical power transfer of the electric machine along a rotation speed range of the second rotational structure by controlling an electrical loading of the electric machine with the control unit.

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.

A hybrid electric turboprop (140, 240, 340) includes a propeller (344), a thermal engine (362), electric motor (370), a gearbox (380, 1380), and an air inlet (348, 1348, 1548) defining an opening (349, 1549). The propeller is configured to rotate about a propeller axis ("P-P") and includes a plurality of blades (354) extending radially outwardly from the propeller axis. As the propeller is rotated, the plurality of blades define a substantially circular blade path area ("A") centered about the propeller axis and having a radius ("r"). The gearbox (380, 1380) couples both the thermal engine and the electric motor to the propeller to enable each of the thermal engine and the electric motor to drive rotation of the propeller about the propeller axis. A radial distance ("i") equal to or less than about 60% of the radius is defined from the propeller axis to a location within the opening.