877results 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.

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.

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.

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.

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.

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.

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 propulsion system for an aircraft includes an engine assembly and a turbocompressor. The engine assembly includes an engine and an interburner. The engine includes an engine output shaft. The engine is configured to drive rotation of a propulsor with the engine output shaft. The interburner is configured to mix and burn an exhaust gas from the engine with fuel to form a combustion gas. The turbocompressor includes a turbine and a compressor. The turbine and the compressor form a rotational assembly. The rotational assembly includes a shaft, a bladed turbine rotor of the turbine, and a bladed compressor rotor of the compressor. The turbine is connected in fluid communication with the interburner to receive the combustion gas. The compressor is connected in fluid communication with the engine to direct a compressed air to the engine. The rotational assembly is mechanically independent of the engine output shaft.

A power system including a power converter system and an electric machine is provided. In one aspect, the power converter system has first and second switching elements. The electric machine includes a first multiphase winding electrically coupled with the first switching elements and a second multiphase winding electrically coupled with the second switching elements. The first and second multiphase windings are arranged and configured to operate electrically opposite in phase with respect to one another. One or more processors control the first switching elements to generate first pulse width modulated (PWM) signals based on received voltage commands to render a first common mode signal and also control the second switching elements to generate second PWM signals based on received voltage commands to render a second common mode signal. The rendered first and second common mode signals have the same or similar waveform with opposite polarity with respect to one another.

A power system including a power converter system and an electric machine is provided. In one aspect, the power converter system has first and second switching elements. The electric machine includes a first multiphase winding electrically coupled with the first switching elements and a second multiphase winding electrically coupled with the second switching elements. The first and second multiphase windings are arranged and configured to operate electrically opposite in phase with respect to one another. One or more processors control the first switching elements to generate first pulse width modulated (PWM) signals based on received voltage commands to render a first common mode signal and also control the second switching elements to generate second PWM signals based on received voltage commands to render a second common mode signal. The rendered first and second common mode signals have the same or similar waveform with opposite polarity with respect to one another.

An Unmanned Aerial Vehicle power management system is configured to match and manage power demands with combined power output from the power storage devices and dynamic generation via a turbine (combustion) motor, while preventing overcurrent and under voltage loads on the power storage devices. The powerplant and power management system of the present invention are, in one embodiment, directed for use with a vertical take-off and landing (VTOL) UAV having multiple forward propulsion propellers (FP) and multiple vertical lift (VL) propellers.

A electrical power system comprises: an electrical machine operable to output AC; a DC electrical network; a power electronics converter connected between the AC output of the electrical machine and the DC electrical network and including a plurality of transistors and associated diodes connected in parallel with the transistors; and a controller configured to control switching of the transistors of the converter so that, during normal operation of the electrical power system, the converter rectifies the AC output of the electrical machine to supply the DC electrical network with DC electrical power. The controller is further configured, responsive to a determination to the effect there is a fault in the DC electrical network, to control a voltage source, to inject a voltage to bias the diodes of the converter, and to control the switching of the transistors to control a level of current supplied to the faulted DC electrical network.

A monitor is configured to monitor a state of a battery mounted on an eVTOL. The monitor includes: an acquisition unit configured to acquire information regarding the battery; a calculation unit configured to calculate, based on the information, a degree of temporary deterioration caused by an imbalance in ion concentration of the battery; and an output unit configured to output information regarding the degree of temporary deterioration. According to the monitor, since the degree of temporary deterioration can be grasped, the flight safety can be enhanced.

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.

The present disclosure relates to a reconfigurable battery system and method of operating the same. An example apparatus includes at least one memory, instructions in the apparatus, and processor circuitry to execute the instructions to determine a state of charge of a battery, determine a closed circuit voltage of the battery, determine a value of a parameter based on a ratio of the state of charge and the closed circuit voltage, and control a switch coupled to the battery based on the value of the parameter, the controlling of the switch to either cause the battery to be coupled to a battery string or cause the battery to be disconnected from the battery string.

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.

A method for managing a power transition between a first power generation mode and a second power generation mode in a turbine engine, the turbine engine having a high-pressure shaft from which high-pressure power is extracted and a low-pressure shaft from which low-pressure power is extracted, the first power generation mode corresponding to a first power distribution between the high-pressure power and the low-pressure power, the second power generation mode corresponds to a second power distribution between the high-voltage power and the low-voltage power, the sum of the high-voltage power and the low-voltage power extracted in each power generation mode defines the same total power, and a transition between the first power generation mode and the second power generation mode occurs within a predetermined time period (t2-t1), meanwhile, the same smoothing function f (t) is applied to the high-voltage power and the low-voltage power, and the total power is not changed.

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.

A modular battery system includes an array of battery modules arranged in at least one stack. Each battery module includes a plurality of battery cells, a first side having positive and negative receptacles and a second side, that is opposite of the first side, having positive and negative plugs. The receptacles and plugs are configured such that adjacent battery modules in a side-by-side relationship are electrically coupled together via plug and receptacle connections and such that the battery modules are electrically coupled together in parallel. An interconnection electrically couples each stack of battery modules together via plug and receptacle connections with one of the battery modules in each stack such that the stacks of battery modules are electrically coupled together in parallel. Each of the battery modules includes a voltage regulator configured to convert voltage between a battery cell voltage and a bus voltage.