401results about "Electric power distribution" patented technology

Aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles. In some embodiments, a system of an aircraft is disclosed, configured to determine one or more desired commands for the aircraft, retrieve engine information for at least one electric propulsion unit (EPU) of a plurality of EPUs of the aircraft, wherein the engine information includes at least one time-based metric for temperature associated with the at least one EPU, generates control commands based on the received engine information, and controls effectors according to the generated control commands to meet the desired commands of the aircraft.

An aircraft wing having a fixed wing and a moveably mounted wing tip device is disclosed. A first cable harness and a second cable harness which are received in a protective sleeve extend between the fixed wing and the wing tip device. The protective sleeve is in a secured configuration in which at least two opposing parts of the protective sleeve are secured together between the first cable harness and the second cable harness to provide the sleeve with a first channel in which the first cable harness is received a second channel in which the second cable harness is received, and a webbing region extending between the first channel and the second channel such that the first and second cable harnesses are held apart by the webbing region.

A wing assembly for an aircraft, the wing assembly having a fixed wing and a wing tip device moveably mounted at the end of the fixed wing. A wiring harness extends between the fixed wing and the wing tip device, a length of the wiring harness being encased in a length of protective sleeve. The protective sleeve is pre-shaped, for example pre-molded, to follow a path along a helical curve.

This disclosure relates to systems and methods for providing a copilot replacement system (CPRS) that enables dual-pilot or multi-pilot aircraft to be operated by a single onboard pilot. Amongst other things, the CPRS solutions can include components that autonomously execute functions traditionally performed by an onboard copilot and / or can establish connections with one or more copilot ground base stations (GBSs) that enable ground-based copilots to remotely provide assistance with operating the aircraft.

Disclosed is a wing assembly for an aircraft. The wing assembly comprises a fixed wing, a wing tip device, a wiring harness and a harness guide. The wing tip device is moveably mounted at a joint at an end of the fixed wing. The wing tip device is moveable about the joint between: (i) a flight configuration for use during flight, and (ii) a ground configuration for use during ground-based operations, in which ground configuration the wing tip device is moved relative to the fixed wing such that the span of the wing is reduced. The wiring harness extends between the fixed wing and the wing tip device. The wiring harness comprises a plurality of conductors and is arranged to transmit electrical power and / or data to the wing tip device. The harness guide is located between the fixed wing and the wing tip device. The harness guide comprises a helical channel configured to receive the wiring harness and guide the wiring harness between the fixed wing and the wing tip device.

Systems, apparatus and methods are provided for securing firmware and mission data in memory of an unmanned aerial vehicle. The unmanned aerial vehicle may include a central body and at least one electric motor. The central body may include a flight controller with a processor, a firmware memory configured to receive and store firmware, and circuit board with a fixed portion and a removable portion with a programming interface connector providing access through a signal connection to a firmware memory operatively connected to the flight controller. The central body may further include a volatile mission memory configured to receive mission data and automatically be erased upon power interruption.

An aircraft propulsion system includes a hybrid-electric power plant for delivering power to an air mover for propelling an aircraft. The hybrid-electric power plant includes a heat engine operatively connected to a first air mover, and an electric motor operatively connected to a second air mover. The second air mover is positioned on a wing of the aircraft outboard from the heat engine. A method for reducing trailing vortices includes powering a first air mover of an aircraft with a heat engine during a take-off stage, a climb stage, a cruise-stage and / or a descent stage. The method includes powering a second air mover of the aircraft with an electrical motor during the take-off stage and / or the climb stage. The method includes freewheeling the second air mover during the cruise stage and / or the descent stage to generate mechanical energy and reduce wing tip vortices.

An aircraft hybrid electrical propulsion (HEP) system includes an electrical system configured to deliver power to a plurality of electrical loads, a propulsion system configured to generate thrust in response to an input power, and an HEP controller in signal communication with the electrical system and the propulsion system. The HEP controller is configured to monitor a load demand of at least one electrical load among the plurality of electrical loads and to actively control the input power to actively control the thrust in response to changes in the load demand.

To provide a power supply system, an aircraft, and a power supply system control method that makes the remaining capacity of each battery approximately equal.SOLUTION: A power supply system 26 includes a power converter 36 that converts the power of any one of a plurality of batteries included in a high voltage circuit 32 and supplies it to one or more second loads 38, a switch 34 capable of selectively connecting any one of the plurality of batteries to the power converter, and a controller 70 that compares the remaining capacities SOC of the batteries, and controls the switch so as to connect a battery with the highest remaining capacity and the power converter when the difference between the highest remaining capacity and the lowest remaining capacity is greater than a predetermined threshold.SELECTED DRAWING: Figure 3

Aspects of the present disclosure generally relate to systems and methods for flight control of aircraft driven by an electric propulsion system and in other types of vehicles. In some embodiments, an aircraft is disclosed, the aircraft comprising: at least one electric propulsion unit; at least one sensor configured to measure at least one aircraft condition; and at least one flight control computer configured to dynamically vary at least one torque command for the at least one electric propulsion unit based at least on the at least one aircraft condition; wherein the at least one electric propulsion unit is configured to generate thrust based on at least one dynamically changed torque command.

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 power distribution circuit for an electrically powered aircraft includes a plurality of batteries and a plurality of electric propulsion systems. A plurality of power distribution circuits each couple a battery of the plurality of batteries to two or more electric propulsion systems. The plurality of electric propulsion systems are positioned on the aircraft to apply balanced forces to the aircraft such that in the event of a failure, the aircraft remains stable and only experiences a loss in altitude or speed.

A power management system for supplying power from an aircraft electrical power source to PEDs. The power management system includes a plurality of power control circuits connected to power outputs configured to supply power to PEDs, and a power management circuit. The power management circuit obtains passenger-specific travel information associated with a passenger corresponding to a respective PED, and determines, based at least in part on the passenger-specific travel information, an allowed subscription power level for the respective PED. The power management circuit controls the power supplied to the respective PED via the respective power control circuit based on the determined allowed subscription power level. The operation to control the power supplied to the respective PED may include to negotiate with the respective PED an agreement for an agreed subscription power level that the respective PED is authorized to be supplied based on the determined allowed subscription power level.

A power distribution system for an aircraft, comprising a plurality of electric propeller units (EPUs), includes a first paired battery pack unit comprising a first battery electrically connected to a second battery via a first high voltage bus. The first and second batteries are configured to provide power to respectively first and second sets of EPUs of the plurality of EPUs. The system includes a second paired battery pack unit comprising a third battery electrically connected to a fourth battery via a second high voltage bus. The third and fourth batteries are configured to provide power to respectively third and fourth sets of EPUs of the plurality of EPUs. The first high voltage bus and the second high voltage bus are electrically separate from one another.

A propulsion system (20) includes a propulsor (26), an engine (22), an electrical assembly (24), and an emergency energy protection assembly. The engine (22) includes a fuel system (38). The fuel system (38) includes a fuel control unit. The electrical assembly (24) includes a battery (66), a motor control unit (64), an electric motor (62), and an electrical distribution system (68). The electrical distribution system (68) includes at least one contactor electrically connected between the battery (66) and the motor control unit (64). The at least one contactor is switchable between a closed state and an open state. The engine (22) and the electric motor (62) are coupled with the propulsor (26). The emergency energy protection assembly includes an emergency actuator. The emergency actuator is actuable in an actuated state to control the at least one contactor to switch to the open state, control the motor control unit (64) to deenergize the electric motor (62), and control the fuel control unit to stop fuel flow for the engine (22).

An aircraft power generation system includes: a turbo fan engine including a variable stator vane which is located upstream of a fan and whose installation angle is variable and a vane actuator that drives the variable stator vane; a power generator mechanically connected to a low-pressure shaft of the engine; and processing circuitry configured to control the vane actuator. The processing circuitry is configured to, when it is determined that the power generator generates electric power during an operation of the turbo fan engine, execute an electric power generating mode of controlling the vane actuator to change the installation angle of the variable stator vane such that the flow rate of air flowing through the fan decreases.

The invention relates to an aircraft, a computer program product, a storage medium, an electrical system for controlling a hybrid aircraft and a method for controlling an electrical system of a hybrid aircraft, the electrical system comprising at least two busbars configured to transfer electrical energy between a plurality of subsystems of the aircraft, the electrical system includes at least one centralized regulation control device, the method including regulation of the voltage level performed by the controller from one or more variable regulation points, the one or more variable adjustment points are determined from a plurality of predetermined adjustment points according to a property of power transmission to be performed and each variable adjustment point is located in any one of the busbars.

An electric propulsion system for an aircraft, comprising: - an electrical power source (1) including a synchronous generator (11) driven by a heat engine (12), - a first synchronous electric motor (3) connected to an output of the synchronous generator by an alternating current electrical circuit (2), - a rotating propulsion unit (7), characterized in that it comprises: - a direct current electrical circuit (4) comprising, from an input connected to the alternating current electrical circuit: a first alternating / direct current converter (41), a direct current electrical link (42) and a second direct / alternating converter (43), - a second electric motor (5) connected to an output of the direct / alternating converter, - a differential gear train (6) including a first input shaft (61) connected to the first synchronous electric motor,a second input shaft (62) connected to the second electric motor and an output shaft (63) connected to the rotating propulsion unit. Figure for the abbreviation: Figure 1,

A vehicle includes a gas turbine engine having at least two spools and an associated power system. The power system includes two independent power subsystems, including a first power subsystem for managing power transfer between spools and a second power subsystem for supplying a base power load to the vehicle. The first power subsystem has a first electric machine mechanically coupled with a first spool of the gas turbine engine and a second electric machine mechanically coupled with a second spool. The second electric machine is electrically coupled with the first electric machine such that electrical power is transmittable therebetween. The second power subsystem has a third electric machine mechanically coupled with one of the spools. The third electric machine is electrically coupled with a load positioned offboard the gas turbine engine. The first power subsystem and the second power subsystem are electrically decoupled from one another.

In an inverter device (80), an inverter high-voltage section (86) is housed in an inverter case (90). The inverter housing (90) has an inverter opening part (92) and an inverter cover part (95). The inverter cover section (95) covers the inverter opening section (92) so as to hide the inverter high-voltage section (86). When the control mode of the flight control device (40) is set to an interlocking mode, power supply from the battery (31) to the motor (61) is cut off as the inverter cover (95) opens. When the control mode is set to the interlock release mode, power supply from the battery (31) to the motor (61) is not cut off even if the inverter cover portion (95) is open. When the eVTOL (10) is in flight, the control mode is set to an interlock release mode.

Electrical power supply system of an aircraft electric propulsion system. The electrical power supply system (10) of an aircraft electric propulsion system (4) (1) comprises: - at least two first fuel cell assemblies (FC11, FC12), intended to electrically power an electric motor driving a propulsion propeller; - a first compressor (C1) configured to supply compressed air to at least some of the at least two first fuel cell assemblies (FC11, FC12); and - a first electric motor (M1) intended to drive the first compressor. The first electric motor (M1) comprises at least two electrical windings (W1a, W1b), each electrical winding being electrically powered from a separate fuel cell assembly among the at least two first fuel cell assemblies (FC11, FC12). Figure for the abstract: Fig. 2

A power distribution system and method of operation for an aircraft includes determining, by a supplemental power system, an operation mode of the aircraft, selecting, by the supplemental power system, at least one supplemental power source of a set of at least two supplemental power sources, and at least one of receiving a first current from the main power bus by the selected at least one supplemental power source, or selectively providing a second current to the main power bus from the selected at least one supplemental power source, based on an operating mode of the aircraft.

A propulsion system (20) for an aircraft (1000) includes a battery (64), an electrical distribution system (66), a ground-based charger, and a battery management system (68). The battery includes a plurality of battery strings (72) each including a plurality of battery cells (76). The electrical distribution system includes a battery string switch assembly (98) operable to electrically interconnect each of the plurality of battery strings together in parallel. The ground-based charger includes a charger controller (132) including a simple electronic hardware (SEH) system. The battery management system includes a battery management system (BMS) controller (108) connected in signal communication with the SEH system. The BMS controller is configured to determine a battery charging profile specific to the battery and charge the battery by controlling the ground-based charger, through the SEH system, to supply electrical power to the electrical distribution system at a set voltage and a set current defined by the battery charging profile.

A propulsion system (20) for an aircraft (1000) includes a battery (64), an electrical distribution system (66), a first control channel (110, 114), and a second control channel (112, 116). The battery (64) includes a plurality of battery strings (72). The electrical distribution system (66) includes at least one string contactor (132, 142) for each of the plurality of battery strings (72). The at least one string contactor (132, 142) includes a first gate (138, 148) connected to the first control channel (110, 114) and a second gate (140, 150) connected to the second control channel (112, 116). The at least one string contactor (132, 142) is operable to a closed position with a first control signal present at the first gate (138, 148) and a second control signal present at the second gate (140, 150). The at least one string contactor (132, 142) is operable to an open position with one or both of the first control signal absent at the first gate (138, 148) or the second control signal absent at the second gate (140, 150).

A power source assembly includes a plurality of strings electrically connectable to a load. Each of the plurality of strings includes a fuel cell module and a DC / DC converter electrically connected to the fuel cell module. Each of the plurality of strings is arranged in parallel to each other of the plurality of strings.

The invention relates to a system (1) for supplying electrical power to at least one aircraft-propelling electric motor (M), the system (1) for supplying electrical power comprising a fuel cell (2), which is configured to be fed with a flow of air (FA) and with a flow of hydrogen (FH) and configured to supply electric power to an electrical distribution unit (3) that is configured to be connected to said propelling electric motor (M), the fuel cell (2) comprising at least a first electrochemical line (LA) and a second electrochemical line (LB) that are mounted electrically in parallel, each electrochemical line (LA, LB) comprising at least two independent stacks (E1-E8) mounted electrically in series, each stack (E1-E8) comprising a plurality of electrochemical cells.

A propulsion system (20) for an aircraft includes a battery, an electrical distribution system (66), and a ground-based charger (86). The battery (64) includes a plurality of battery strings (72). Each of the plurality of battery strings (72) includes a plurality of battery cells (76). The electrical distribution system (66) includes a battery string switch assembly (98). The battery string switch assembly (98) is operable to electrically interconnect each of the plurality of battery strings (72) together in parallel. The ground-based charger (86) is electrically connected to the electrical distribution system (66). The ground-based charger (86) includes a charger controller (132). The charger controller (132) is configured to determine a battery charging profile (138) specific to the battery (64) and charge the battery (64) by supplying electrical power to the electrical distribution system (66) at a set voltage and a set current defined by the battery charging profile (138).

In a case that supply of DC power from a first power generation device (14a) to a first power supply circuit (18a) is disrupted, a control device (30) executes, on a first inverter (15a), a first power suppression control to reduce electric power supplied from the first inverter to a first motor (17a) to be smaller than the electric power prior to the supply of DC power from the first power generation device to the first power supply circuit being disrupted, and also executes, on a second inverter (15b), a first power increasing control to increase electric power supplied from the second inverter to the first motor to be larger than the electric power prior to the supply of DC power from the first power generation device to the first power supply circuit being disrupted.

An aircraft power supply network and associated aircraft, the power supply network includes a first engine and a second engine, a first and a second set of loads, an auxiliary electrical source and a first electrical machine connected to the first engine. The power supply network also includes a first converter connected to the first electrical machine, a first interconnection bus connected to the first converter by a first primary switch and to the first set of loads and a second electrical machine connected to the second motor. The power supply network further includes a second converter, a second interconnection bus, connected to the second converter by a second primary switch and to the second set of loads, a sharing bus connected to the first interconnection bus by a first inter-bus switch, connected to the second interconnection bus by a second inter-bus switch and connected to the auxiliary electrical source.