125results about "Depending on power plant type" patented technology

A reduction gearbox of an aircraft propulsion system includes an input shaft configured to input rotational energy into the reduction gearbox, an output shaft configured to output rotational energy from the reduction gearbox to a propulsor of the aircraft propulsion system, and an epicyclic gear arrangement mechanically coupling the output shaft to the input shaft. A transfer gear is positioned at a component of the epicyclic gear arrangement and is configured to mechanically couple an accessory component to the reduction gearbox.

A rotor assembly includes a propeller, a motor, and a connection assembly. The motor includes a stator and a rotator rotatable with respect to the stator, and is configured to drive the propeller to rotate through the rotator. The connection assembly is arranged between the propeller and the motor, and includes a locking member and an elastic member. The elastic member is configured to provide a force onto the locking member in response to the propeller being locked to the motor by the locking member, to maintain a locking state of the locking member.

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).

The invention relates to an electric propulsion system comprising: - a propeller (32) having an axis of rotation (A32), - at least two electric motors (34) offset from each other along a longitudinal direction parallel to the axis of rotation (A32) of the propeller (32), - a main drive shaft (42) parallel to the axis of rotation (A32) of the propeller (32) to which the electric motors (34) are coupled, - a transmission system (36) comprising an output shaft (38) connected to the propeller (32) and an input shaft (38') connected to the main drive shaft (42). This solution simplifies the design of the transmission system and reduces the cross-section of the electric propulsion system at the location of the electric motors. The invention also relates to an aircraft comprising at least one such electric propulsion system.

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.

An aircraft system (20) includes a first propulsor (36A), a second propulsor (36B), a drivetrain and an intermittent combustion engine (66). The first propulsor (36A) includes a first propulsor rotor (42A) and a first vane array (44A). The second propulsor (36B) includes a second propulsor rotor (42B) and a second vane array (44). The drivetrain includes a drive structure (100) and a transmission (106). An output (118) of the transmission (106) is coupled to the first propulsor rotor (42A) and the second propulsor rotor (42B) through the drive structure (100). The intermittent combustion engine (66) is configured to drive rotation of the first propulsor rotor (42A) and the second propulsor rotor (42B) through the drivetrain.

The invention relates to a monitoring system for monitoring a hybrid power plant of a rotorcraft, an associated rotorcraft and a method. The hybrid power plant (2) comprises: an electric motor (5) comprising a first drive shaft (15); comprising a second drive shaft (18); and a gearbox (9). The monitoring system comprises: at least one first sensing device (16, 17, 19) measuring a first item of torque information representative of a first engine torque C1 transmitted by a first drive shaft (15) to a first input shaft (10) of a gearbox (9); and a controller (31) that compares the first engine torque C1 with a first limit value VL1.

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.

In examples, provided are leadless power couplers that include (1) a thermal insulating system having an outer wall and an inner wall, (2) a first electrically conductive winding located outside the thermal insulating system, where the first electrically conductive winding is configured to create a varying magnetic field, (3) a plurality of second electrically conductive windings located inside the thermal insulating system and configured to couple to the varying magnetic field, the plurality of second electrically conductive windings being superconductors, (4) a plurality of cryogenic rectifiers, each cryogenic rectifier being coupled to a respective second electrically conductive winding in the plurality of second electrically conductive windings, and (5) a plurality of cryogenic cables coupled between respective outputs of the plurality of cryogenic rectifiers and respective loads.

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 present disclosure provides a propulsion assembly for mounting on an aircraft platform, the propulsion assembly configured to generate thrust and enable controllable accelerated inflow and outflow of a mass flow of air. The propulsion assembly includes one or more fluid-based rotary actuators or motors in fluid communication with a source of active / pressurized working fluid, driving the one or more fluid-based rotary actuators by receiving the active working fluid. The propulsion assembly further comprises one or more actuator disk devices, each including a plurality of blades rotatable by one or more fluid-based rotary actuators. The valve mechanism of the propulsion assembly is disposed along a flow path between the source of active working fluid and the one or more fluid-based rotary actuators. The propulsion assembly further comprises at least one processing circuit and one or more memories connected to the at least one processing circuit and storing programming instructions executed by the at least one processing circuit, (1) the at least one processing circuit executing the programming instructions to control the valve mechanism so as to effect at least one of (i) a selected flow rate of the active working fluid within the one or more fluid-based rotary actuators or the active working fluid received by the one or more fluid-based rotary actuators, (ii) a selected fluid pressure, and (iii) a selected temperature. This configuration of the propulsion assembly enables control of the instantaneous angular acceleration and rotational speed of the motor, and thus enables control of the rotational movement of the actuator disk device.

The present invention relates to a monitoring system (1) for monitoring a hybrid power plant (2) of a rotorcraft (3) comprising at least one rotor (4), said hybrid power plant (2) comprising an electric motor (5) comprising a first drive shaft (15), a heat engine (8) comprising a second drive shaft (18) and a power transmission box (9). The monitoring system comprising at least a first sensor (16, 17, 19) measuring first torque information representative of a first engine torque C1 transmitted by the first drive shaft (15) to a first input shaft (10) of the power transmission box (9) and a controller (31) comparing said first engine torque C1 with a first limit value VL1. Abstract figure: figure 1

The application discloses an oil-electric hybrid unmanned aerial vehicle power system and relates to the technical field of unmanned aerial vehicles and aircrafts. The application is characterized by the coaxial direct-drive direct-connection cooperation design of the propeller, the engine and the direct-current permanent magnet motor, simple structure, reliability, durability and high power-to-weight ratio. The oil-electric hybrid system is formed by the cooperation of the oil engine and the permanent magnet direct-current motor. The permanent magnet direct-current motor can start the engine and convert the excess power into electric energy to charge the storage battery. The torque of the engine can be changed by the direct-current bidirectional electric governor, so the speed of the propeller can be accurately controlled, and a three-in-one machine is formed. Different displacement engines can be matched with corresponding power direct-current permanent magnet motors. Different power ratios can be used for different models of unmanned aerial vehicles and multi-rotor aircrafts. Different accurate control and speed regulation ratios can be combined by the cooperation of different power engines and direct-current permanent magnet motors.

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.

To control charging of a battery without applying an overload by power feed of a generator based on a state of charge of the battery. A charge control system includes an engine and a motor generator for supplying generated power to two motors of a VTOL rotor and a cruising rotor, a plurality of batteries including two batteries for storing the electrical power generated by the generator and supply the stored electrical power to each of the two motors, a switch for connecting or disconnecting each of the batteries to or from the generator, and a control section for controlling the switch based on a state of charge of each of the batteries to disconnect one battery with a larger charge amount out of the two batteries from the generator and connect the other battery with the smaller charge amount out of the two batteries to the generator.

The utility model discloses a power system of an aircraft and a multi-rotor aircraft based on hybrid power. The power system comprises a fuselage main body, the main rotor assembly is powered by fuel oil; the power system comprises a controller; the auxiliary rotor assembly is powered by the battery pack; a transmission assembly; and a seat. According to the utility model, when one engine of the aircraft breaks down, the aircraft can still work; through the dynamic power distribution technology of the controller, the aircraft can save electric energy, and energy waste is reduced; fuel oil and a battery pack are used for providing power, so that the multi-rotor aircraft has the performances of large load, sufficient power, stable flight attitude and the like; and the engine for providing kinetic energy for the main rotor assembly can charge the battery pack of the auxiliary rotor assembly, so that the cruising ability and the practical efficiency of the multi-rotor aircraft are remarkably improved.

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.

The present invention relates to a method for controlling a thermal and electrical power plant (5) for setting in motion at least one rotary member (2, 3) of a rotorcraft (1), said power plant (5) comprising at least one thermal engine (11) and an electrical system (20) provided with at least one electrical machine (21). The method comprises: - selection with a selector (42) of an operating mode chosen from several operating modes, - determination of a density altitude and comparison with a controller (25) of said current density altitude and a threshold density altitude, - control with the controller (25) of said at least one electrical machine (21) as a function at least of said chosen operating mode as well as of said comparison and of a power required to be supplied to said power transmission chain (15). Abstract figure: figure 1

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).