1188results 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 computer-implemented method for controlling an aircraft is disclosed, comprising: receiving, from a source processor, a first copy of a signal corresponding to an input device; sending a second copy of the signal to all other processors; receiving a number of second copies of the signal from all other processors, the number of second copies being equal to the number of all other processors excluding the source processor; determining a consensus signal based on the first copy and the second copies of the signal; and determining a command signal for an effector of the aircraft based on the consensus signal, and wherein no two processors are configured to receive signals from a same input device.

An electrical propulsion system for a vertical take-off and landing (VTOL) aircraft comprises an electrical motor assembly and an inverter assembly. The inverter assembly comprises a housing, a capacitor assembly, at least one printed circuit board assembly (PCBA), and a plurality of positioning pins. The capacitor assembly comprises a center hole, at least one capacitor, a capacitor housing having at least one busbar, and a plurality of through holes in the capacitor housing. The capacitor assembly and the at least one PCBA are positioned inside the housing. The plurality of positioning pins pass through the through the plurality of through holes of the capacitor housing and the at least one PCBA and are connected to the housing.

A power source assembly is provided having a fuel cell module configured to provide a first direct current power output; a battery module configured to provide a second direct current power output; a direct current electric bus configured to provide a net direct current power output to a load; a DC / DC converter in electrical connection with the direct current electric bus, the DC / DC converter configured to receive the first direct current power output from the fuel cell module or the second direct current power output from the battery module; and a controller operably coupled to the DC / DC converter and configured to receive data indicative of the first direct current power output, the controller configured to control the DC / DC converter based on the data indicative of the first direct current power output to maintain a slew rate from the fuel cell module within a slew rate range.

A vertical take-off or landing (VTOL) aircraft is disclosed that includes a propulsion system comprising at least four tilting propulsion assemblies and two or more lifting propulsion assemblies. Actuation of the tilting mechanisms can convert the aircraft from a hover configuration, in which all propellers produce thrust generally upward to counteract the aircraft's weight, to a forward flight configuration, in which the four or more tilting propellers produce thrust generally in a forward direction parallel to the fuselage to overcome drag in cruise flight and the lifting propellers produce no thrust or a level of thrust generally less than that required in the hover configuration.

The present invention relates to a system for converting and transporting electrical energy in a turbomachine-propelled aircraft, said system including:at least a first rotary electric machine connected mechanically to the low-pressure shaft by means of a coupling device,a second rotary electric machine connected mechanically to the high-pressure shaft by means of an accessory gearbox, andsaid first electric power module and said second electric power module being connected electrically to an internal electrical network of the turbomachine,a generator coupled to the accessory gearbox and intended to supply an electrical network of the aircraft with electricity, anda control device for controlling start-up of the turbomachine by at least one rotary electric machine, for compensating for power consumed by the generator, and for implementing distribution between the power absorbed by the high-pressure shaft and the power absorbed by the low-pressure shaft.

A propulsion system, comprising: a motor assembly including a stator and a rotor, a heat exchanger configured to receive air to cool a liquid, and a thermal plate including one or more channels configured to convey the liquid from the heat exchanger to the motor assembly. The heat exchanger is mechanically and fluidically coupled to the thermal plate.

An aircraft electrical power supply system includes a fuel cell auxiliary power unit (APU) that supplies auxiliary electrical power to an aircraft, a fuel cell power plant that supplies primary electrical power to the aircraft, and a hydrogen storage unit that supplies hydrogen to the fuel cell APU and the fuel cell power plant.

A control system for charging an aircraft, comprising: a battery pack, an input device configured to enable a user to select between different charging modes, two main contactors connecting the battery pack to an electric propulsion unit (EPU) load and an auxiliary load, a EPU load contactor connecting the battery pack to the EPU load and a controller configured to receive the selected charge mode and control the contactors, keep the two main contactors open upon receiving a user selection to charge in a first mode, close the two main contactors and keep an EPU load contactor open upon receiving a user selection to charge in a second mode, and close the two main contactors and the EPU load contactor upon receiving a user selection to charge in a third mode.

A method of determining the capability of a nonlinear aircraft power source includes accessing status values representing a current state of an energy storage system in an aircraft, accessing demand values related to expected power demands on the energy storage system, modeling an ongoing status of the energy storage system using the status values and the demand values to predict when one of the status values will reach a threshold value, and providing an output of a capability of the aircraft based on the status value reaching the threshold value.

The present disclosure relates generally to controlling an aircraft to avoid overheating components. In one embodiment, a method is disclosed, comprising: receiving first sensor data indicating at least one attribute of high voltage wiring of the aircraft; receiving a state of operation of the aircraft; determining a proximity of a temperature of the high voltage wiring to a temperature limit based on the first sensor data and the state of operation of the aircraft; and controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit based on the determined proximity.

The invention discloses an electric aircraft, a power distribution system, a power supply control method and a computer storage medium, and the power distribution system of the electric aircraft comprises a first battery pack, a second battery pack, a first isolation device and a second isolation device. Wherein the first battery pack and the second battery pack supply power to the first part of rotor wing motors through the first isolation device, supply power to the second part of rotor wing motors through the second isolation device, and the first isolation device and the second isolation device perform isolation and power reconstruction of a fault line when the power distribution system has a power supply fault. The embodiment of the invention is simple in structure and low in cost.

A power reduction system for an energy storage system of an aircraft includes a controller configured to control power reduction of power supplied from the energy storage system to an aircraft engine supply bus; and an override switch configurable in an override state and a non-override state. The override switch is configured to: in the non-override state, permit the controller to control the power reduction according to a default configuration comprising one or more parameters that trigger the power reduction; and in the override state, control the power reduction to be performed according to a relaxed configuration that at least one of relaxes and omits the one or more parameters in the default configuration.

An electrical propulsion system comprises an inverter, an electrical motor assembly, an assembly, and a rotor position sensor. The inverter comprises a printed circuit board assembly (PCBA). The electrical motor assembly comprises a stator and a rotor. The assembly is configured to rotate a propeller and comprises a moving component and a stationary component. The rotor position sensor comprises at least one sensor coupled to the PCBA, and a magnet located on the moving component. The at least one sensor is configured to detect a magnetic field of the magnet through the stationary component.

A system for an aircraft includes a fan rotor, an engine core, a fan case, a core case, a bifurcation, an inner electrical system, an outer electrical system and multiple electrical lines electrically coupling the inner electrical system to the outer electrical system. The engine core is configured to drive rotation of the fan rotor about an axis. The fan case houses the fan rotor. The core case houses the engine core. The bifurcation radially crosses a bypass flowpath. The bypass flowpath is downstream of the fan rotor and located radially between the fan case and the core case. The inner electrical system is arranged with the engine core. The outer electrical system is mounted to the fan case. The electrical lines extend between and electrically couple the inner electrical system to the outer electrical system. The electrical lines run in parallel along one another radially through the bifurcation.

A electric aircraft power distribution system includes a first battery pack connected to at least a first load and to a common bus that connects the first battery pack in parallel to at least a second battery pack; a first electrical component electrically connected between the first battery pack and the first load and configured to disconnect the first load from the first battery pack in response to current above a first threshold current, wherein the first electrical component has a first disconnection time at the first threshold current; and a second electrical component electrically connected between the first battery pack and the common bus and configured to disconnect the first battery pack from the common bus in response to current above a second threshold current, wherein the second electrical component has a second disconnection time at the second threshold current that is higher than the first disconnection time.

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 with a folding wing tip arrangement having a wing tip device mounted at a joint at the end of a fixed wing. It is desirable to transfer power and / or data into a folding wing tip, which presents challenges such as repeated exposure to potentially harsh environmental conditions, with movement, and / or with changes in tensional loads. A wiring harness includes a plurality of conductors and is arranged to transmit electrical power and / or data to the wing tip device. A length of the harness is received in a protective sleeve, and the sleeve extends over a length of the wiring harness spanning the joint, the sleeve being removable from the wiring harness.

An aircraft with a folding wing tip arrangement and first and second wiring harnesses extending between a fixed wing and a wing tip device is disclosed. It is desirable to transfer power and / or data into a folding wing tip, which presents challenges such as repeated exposure to potentially harsh environmental conditions, with movement, and / or with changes in tensional loads. The wiring harnesses include a plurality of conductors and are arranged to transmit electrical power and / or data to the wing tip device. The sleeve includes two channels to receive the harnesses, the position of the channels being fixed, relative to one another, such that the harnesses are held apart.

A system for dynamic excitation of an energy storage element configured for use in an electric aircraft includes a plurality of propulsors mechanically connected to the electric aircraft. The system includes a plurality of energy storage elements electrically connected to the plurality of propulsors. The system includes a bus element, wherein the bus element is electrically connected to the plurality of energy storage elements and a cross tie element connected to the bus element configured to disconnect a first energy storage element from a second energy storage element. The system includes a modulator unit electrically connected to the bus element configured to modulate a first electrical command to the first energy storage element and a second electrical command to the second energy storage element. The system includes at least a sensor configured to detect an energy datum corresponding to the first energy storage element as a function of the modulation.

An electrical propulsion system for a vertical take-off and landing (VTOL) aircraft comprises an electrical motor assembly and an inverter assembly. The inverter assembly comprises a housing, a capacitor assembly, at least one printed circuit board assembly (PCBA), and a plurality of positioning pins. The capacitor assembly comprises a center hole, at least one capacitor, a capacitor housing having at least one busbar, and a plurality of through holes in the capacitor housing. The capacitor assembly and the at least one PCBA are positioned inside the housing. The plurality of positioning pins pass through the through the plurality of through holes of the capacitor housing and the at least one PCBA and are connected to the housing.

A method for controlling a hybrid power plant of a vehicle provided with a heat engine and at least one electrical unit comprising an electric machine, an electrical energy source, and a power management unit, the heat engine and the at least one electric machine being mechanically connected to input shafts of a gearbox. After acquiring parameters of the at least one electrical unit, and receiving a mechanical power requirement of the vehicle, a usable electrical power that can be used by the electrical unit and at least one operating setpoint of the electrical unit are determined as a function of the parameters and the mechanical power requirement. Finally, the heat engine and the electrical unit are controlled as a function of the at least one operating setpoint in order to meet the mechanical power requirement.

An electric power system (400) for a vehicle includes at least one electric machine (321), one or more power rectifiers, and a plurality of DC channels. The at least one electric machine (321) includes a plurality of tooth-wound multi-phase windings (311) that are substantially magnetically decoupled, and the at least one electric machine (321) is mechanically balanced even if one of the plurality of windings (311) is de-energized. The one or more power rectifiers are configured to produce rectified power from the power generated by the at least one electric machine (321). The plurality of DC channels are formed after the at least one power rectifier and are configured to provide DC power to one or more loads within a vehicle.

A method includes utilizing a learning model, which has been trained to mimic the nominal control outputs of a model predictive control, to generate a nominal control output for a system that is an aircraft system having one or more system constraints. The method also includes, based on the nominal control output, satisfying each of one or more system constraints, implementing the nominal control output; and, based on the nominal control output not satisfying one of the one or more system constraints, utilizing a safety stability filter that includes a control barrier function (CBF) portion and a control Lyapunov function (CLF) portion to modify the nominal control output and obtain a modified control output, and implementing the modified control output in the aircraft system. A system is also disclosed.

An electric propulsion system for a vertical take-off and landing aircraft has an inverter assembly, a gearbox assembly, an electric motor assembly. The inverter assembly, the gearbox assembly, and the electric motor assembly may be substantially aligned along an axis. The inverter assembly, the gearbox assembly, and the electric motor assembly may each abut at least one of the other assemblies. The power engine may use a liquid, such as oil, to cool and lubricate components of the inverter assembly, the gearbox assembly, and the electric motor assembly. Further, the power engine may use a volume of oil below a threshold volume for cooling and lubricating components such that the power engine does not require a fire barrier.

“A battery pack assembly, comprising: a battery pack, the battery pack comprising: a battery pack enclosure, and one or more battery cells within the battery pack enclosure, and a junction box, the junction box comprising: a junction box enclosure including one base wall, four side walls connected to the base wall, and one open end, and a battery management unit, at least one fuse, and at least one switch, each located within the junction box enclosure and mounted to the base wall. The four side walls of the junction box enclosure are removably secured to the battery pack enclosure at the open end.”

Disclosed is a crash handling system for an aircraft, comprising an altitude surveillance system adapted to receive altitude data from an altitude sensor and issue an altitude warning signal when a predetermined altitude condition is fulfilled, a crash detection unit, comprising an acceleration detector and issues a crash detection signal when a predetermined acceleration condition is fulfilled, a control unit operatively coupled to the altitude surveillance system and the crash detection unit to receive respective altitude warning signals and crash detection signals; and an energy storage system isolation means operatively coupled to the control unit to electrically isolate an energy storage system of the aircraft upon reception of an activation signal, wherein the control unit transitions from a disabled operation state to an enabled operation state when the altitude warning signal is received; and to issue the activation signal upon reception of the crash detection signal when in enabled state.

A wireless charging table for use in an aircraft includes a frame assembly for mounting the wireless charging table within the aircraft; a table leaf support coupled to the frame assembly and maneuverable between a stowed position and a deployed position; a main table leaf mounted to the table leaf support and having a power unit; and an electrical system to provide power to the power unit when the table leaf support is in the deployed position. When the table leaf support is in the deployed position, an electrical connection is established between a power source and the power unit, and when the table leaf support is in the stowed position, the electrical connection is broken between the power source and the power unit. The power unit is to provide wireless charging to one or more devices when the electrical connection is established.

The invention relates to a power supply architecture (1) for a hybrid-electric propulsion aircraft (0), comprising a gas turbine (2), a propulsion power supply network (5), and a non-propulsion power supply network (6); the propulsion power supply network (5) comprising a first AC generator (9), at least one first AC / DC converter (10), and at least one first battery (11) for supplying power to the propulsion equipment (7) of the aircraft; the non-propulsion power supply network (6) comprising a second AC generator (20), at least one second AC / DC converter (21), and at least one second battery (22) for supplying power to the non-propulsion equipment (8) of the aircraft.

A system of a hybrid aircraft includes a first gas turbine engine, a second gas turbine engine, and a controller. The first gas turbine engine includes a first low spool electric machine and a first high spool electric machine. The second gas turbine engine includes a second low spool electric machine and a second high spool electric machine. The controller is operable to determine an operating mode of the hybrid aircraft and control power extraction from either or both of the first low spool electric machine and the first high spool electric machine while a single engine descent mode is active. Electric power is provided to either or both of the second low spool electric machine and the second high spool electric machine while the single engine descent mode is active to balance thrust between the first gas turbine engine and the second gas turbine engine.

A power supply system can include an electrical power source and a power converter connectable to a load. The electrical power source can provide a first DC voltage with a first potential difference (VH) between a first DC potential output and a second DC potential output, and a third DC potential output having a potential between the potential at the first DC potential output and the potential at the second DC potential output. The power converter can provide a second voltage with a second potential difference (VL) between a first potential output (V+) and a second potential output (V−). The second potential difference (VL) can be less than the first potential difference (VH). The power converter can electrically isolate the first DC voltage from the second voltage, and the second potential output (V−) can be connected to the third DC potential output using a first connection including an impedance.