583 results about "Electric aircraft" patented technology

Described herein is an electric aircraft comprising a battery pack to power the aircraft. The battery pack comprises a housing, an arrangement of battery cells within the housing and a crumple structure associated with the battery pack. The crumple structure is beneath the arrangement of battery cells and is adapted to compress in the event of an impact from beneath the aircraft. A venting channel is fluidically coupled to the battery cells and arranged to vent gasses away from the battery cells. The venting channel passes through the crumple structure and is adapted to deform with the crumple structure in the event of an impact from beneath the aircraft.

The present disclosure relates generally to flight control of electric aircraft and other powered aerial vehicles. In one embodiment, a method is disclosed, comprising: receiving a descent rate command from a pilot input device, determining a proximity of each propeller of at least two propellers to a vortex ring state; and controlling the aircraft's descent rate to be less than the commanded descent rate when at least one of the at least two propellers is within a first threshold proximity to the vortex ring state.

In an aspect an apparatus for determining a most limiting parameter of an electric aircraft is presented. An apparatus includes at least a processor and a memory communicatively connected to the at least a processor. A memory contains instructions configuring at least a processor to receive aircraft data from a sensing device. A sensing device is configured to measure a parameter of an electric aircraft and generate aircraft data. At least a processor is configured to determine a most limiting parameter of an electric aircraft as a function of aircraft data. At least a processor is configured to communicate a most limiting parameter to a pilot indicator in communication with the at least a processor and memory communicatively connected to the at least a processor. A pilot indicator is configured to display a most limiting parameter to a user.

The invention discloses a thrust compensation control method and system based on a multi-duct electric aircraft, and belongs to the technical field of aviation aircraft control. When the aircraft runs, a motor controller monitors running data of each motor in real time and diagnoses the fault state of each motor in real time, and once a motor triggers power degradation or failure fault, the motor controller can start the motor controller to start the motor controller; and the electric propulsion control unit distributes the thrust of the remaining motors based on an optimization algorithm. Through the above mode, the system can quickly identify the faulty motor and trigger the thrust optimization algorithm in real time by monitoring the running state of the motor in real time, after the fault occurs, the system calculates the current total thrust and total torque according to the initial thrust distribution and the motor position, redistributes the thrust of the remaining non-faulty motors through the optimization algorithm, and the thrust of the non-faulty motors is optimized. The total thrust and the total torque are kept consistent with those before the fault, the attitude of the aircraft is prevented from being out of control, the service life of a motor battery is balanced by minimizing the thrust difference, and the system reliability is improved.

A system for pre-charging short circuit detection in electric aircraft. The system includes a charging connector and a controller communicatively connected to the charging connector. The charging connector includes a housing, at least a conductor, and at least a control signal conductor. The housing is configured to mate with an electric aircraft port of an electric aircraft. The at least a conductor is configured to conduct a charging current to the electric aircraft. The at least a control signal conductor is configured to conduct a control signal. The controller is configured to receive a circuit health datum, through the at least a control signal conductor, prior to initiation of charging of the electric aircraft, and to disable initiation of charging of the electric aircraft if the circuit health datum comprises a short circuit datum. A method for pre-charging short circuit detection in electric aircraft is also provided.

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 system and a method for controlling a propulsor assembly of an electric aircraft. The system may include an electric motor, wherein the electric motor may include a rotor and a stator, a propulsor driven by the electric motor configured to propel an electric aircraft. The propulsor may include a propeller. The propeller may include a blade. The system may include a cyclic control assembly configured to deflect the blade. The cyclic control assembly may include an actuator and a push rod mechanically connected to the actuator and the propulsor. The system may include a flight controller. The flight controller may be configured to receive sensor datum from at least a sensor communicatively connected to the electric aircraft, generate a deflection command as a function of the sensor datum and actuate the cyclic control assembly as a function of the deflection command.

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.

A hybrid-electric aircraft propulsion system includes a gas turbine engine and an electrical assembly. The gas turbine engine includes a first rotational assembly. The first rotational assembly is rotatable about a rotational axis of the gas turbine engine. The first rotational assembly includes a first shaft, a bladed first compressor rotor, and a bladed first turbine rotor. The first shaft interconnects the bladed first compressor rotor and the bladed first turbine rotor. The electrical assembly includes a first motor-generator, a first motor control unit, a motor-generator (MG) cooling system, and a motor control unit (MCU) cooling system. The first motor-generator is coupled to the first shaft. The first motor control unit is electrically connected to the first motor-generator. The MG cooling system is connected in fluid communication with the first motor-generator. The MCU cooling system is connected in fluid communication with the first motor control unit. The MCU cooling system is independent of the MG cooling system.

A system for state determination of a battery module configured for use in an electric vehicle. The system including a battery module including at least a battery cell, a sensor including a proximity sensor configured to detect a status datum corresponding to the battery module, a processor configured to receive the status datum from the sensor, generate a charge datum as a function of the status datum corresponding to the battery module, generate a health datum as a function of the status datum corresponding to the battery module, transmit the charge datum and the health datum, and a display configured to receive the charge datum and the health datum corresponding to the battery cell, and display the charge datum and the health datum corresponding to the battery cell.

A system for electric aircraft fleet management for at least an electric aircraft is provided. The system includes a computing device communicatively connected to at least an electric aircraft, wherein the computing device is configured to receive measured data from the at least an electric aircraft, generate the aircraft performance model output as a function of the measured data, transmit the aircraft performance model output to a user, receive a flight plan from the user; and manage the at least an electric aircraft using the flight plan.

A system including a engine having a low spool and a high spool, a first hybrid electric motor connected to the low spool and a second hybrid electric motor connected to the high spool. An auto-throttle controls an amount of power provided to the engine responsive to at least one aircraft parameter. A power splitting algorithm implemented between the auto-throttle and at least one of the first and second hybrid electric motors determines a power split dividing the total engine power into the power produced by burning fuel and electrical power provided to the at least one of the first and second hybrid electric motors responsive to control signals from the auto-throttle.

A hybrid propulsion update system includes a controller in signal communication with a replaceable battery. The controller includes a tuning parameters storage unit configured to store at least one tuning parameter corresponding to the replaceable battery. The controller is configured to execute at least one optimization algorithm that utilizes the tuning parameters to control operation of the hybrid electric aircraft according to a first performance. The tuning parameters storage unit is configured to receive at least one updated tuning parameter from a controller updating device. The controller executes the at least one optimization algorithm that utilizes the at least one updated tuning parameter such that the hybrid electric aircraft operates according to a second performance that improves upon the first performance.

The invention discloses an electric aircraft power battery management system with a fault isolation function, which comprises a series branch, the series branch comprises a plurality of battery modules connected in series, each battery module is connected in series with an isolation switch, and each battery module is connected in parallel with a bypass switch; the series branch is configured to be used for managing a battery module management unit of the battery module, and the isolating switch and the bypass switch are both controlled by the battery module management unit; the plurality of series branches are connected in parallel and converged on the high-voltage direct-current bus, and the series branches are connected with the high-voltage direct-current bus through the access switch; and the battery module management unit is connected with the main control unit through an internal data bus. The bypass switch circuit is added to the battery module in the battery pack, when the system detects that the voltage of the battery module is abnormal, the battery module can be bypassed, other battery modules can continue to work, and the safety of the battery system can be remarkably improved.

An electric aircraft according to an embodiment of the present invention comprises: a fuselage equipped with a power means, a front spar and a rear spar extending from the fuselage to an end of a wing, and a plurality of ribs extending from the rear spar to the front spar and coupled to the front spar and the rear spar, in which a plurality of solid state batteries are mounted in a plurality of individual spaces partitioned by the front spar, the rear spar, and the plurality of ribs, respectively, and the front spar and the rear spar are used as members for serial connection of the plurality of solid state batteries, and the plurality of ribs are used as members for parallel connection of the plurality of solid state batteries.

In an aspect a system for energy tracking in an electric aircraft. A system includes at least a battery pack. At least a battery pack includes a plurality of battery modules. A system includes a sensing device. A sensing device is configured to detect a battery parameter of at least a battery module of a plurality of battery modules. A sensing device is configured to generate battery data as a function of a detected battery parameter of at least a battery module. A system includes a computing device. A computing device is in electronic communication with a sensing device. A computing device is configured to receive battery data from a sensing device. A computing device is configured to determine an energy amount of a plurality of battery packs as a function of battery data.

The invention discloses an energy management control system and method for a fuel cell electric aircraft, and belongs to the technical field of fuel cells, the energy management control system comprises a fuel cell assembly, a first DC / DC converter, a lithium battery, a solar panel, a second DC / DC converter, an energy management controller and a touch display module, and the lithium battery is used for outputting a second power signal. Energy is distributed according to needs, for example, when the energy consumption of a flight task is high, the hybrid power mode is automatically switched, the energy supply flexibility is enhanced, the touch display module gives consideration to task demand input and energy state display, the operation convenience and the monitoring intuition are improved, finally, the energy utilization efficiency is greatly improved, and the endurance time of the aircraft is prolonged. And meanwhile, the stability and reliability of energy supply in different task scenes are guaranteed.

A system for cross-channel communication for effectors in an electric aircraft is presented. The system may include a flight component of an electric aircraft. The system may include a plurality of effectors, wherein the plurality of effectors may be configured to control the flight component. The system may include a plurality of flight controllers communicatively connected to the plurality of effectors, wherein the plurality of flight controllers may be configured to receive an input, generate a command as a function of the input and transmit the command to the plurality of effectors. The system may include a plurality of networks communicatively connected to the plurality of effectors and the plurality of flight controllers, wherein the plurality of networks may be configured to receive the command from the plurality of flight controllers and transmit the command to the plurality of effectors.

The invention discloses an electric aircraft power load controller system and a working method thereof, and belongs to the technical field of electric aircraft controllers, and the electric aircraft power load controller system comprises a main power module, a pre-charging module, a discharging module, a monitoring control module, a power supply module, a communication module and a terminal. Through the mode, the solid-state power switch and the high-power mechanical switch are connected in parallel, and the control time sequence of the two switches is set, so that the high-voltage high-power electric load circuit can be suitable for the application scene of a high-voltage high-power electric load of an electric aircraft, and has good capacitive load adaptability and electric arc protection performance; and controllable pre-charging of the equivalent capacitance of the high-voltage line is realized by controlling the semiconductor power switch by utilizing the impedance characteristic of the semiconductor power switch in the saturation region through the pre-charging module, so that the complexity and the weight of equipment are reduced.

The present disclosure is directed to systems and methods for cooling an electric aircraft. The system comprises an electronic device, a casing, a fin, and at least one PHP. The casing comprises an inner surface. A PHP may be coupled to the inner surface. The fin comprises a fin interior cavity. A PHP may be embedded within the fin interior cavity. The same PHP may be both coupled to the casing inner surface and embedded within the fin interior cavity, or there may be a plurality of fins such that at least one fin is coupled to the casing inner surface and at least one fin is embedded within the fin interior cavity. The portion of the PHP closest to the electronic device comprises an evaporator section and the portion of the PHP furthest from the electronic device comprises a condenser section.

An electric aircraft includes: one or two or more electric propulsion systems each including a propeller or fan for propulsion disposed to contribute to a lift of a main wing and an electric motor that drives the propeller or fan; and a controller that adjusts the electric propulsion system on the basis of a relationship between a variable relating to an operating state of the electric propulsion system and an aerodynamic force generated on the main wing such that a total thrust by the electric propulsion systems or the aerodynamic force on the main wing has a predetermined value or falls within a predetermined range.

A hybrid electric engine control module (ECU) configured to be operatively connected to a hybrid electric aircraft powerplant having a heat engine system and an electric motor system to control a torque output from each of the heat engine system and the electric motor system. The ECU can be configured to receive a torque command and split output power between the electric motor system and the heat engine system. Additionally and / or alternatively, the ECU can be configured to balance a total torque against a second total torque of a second aircraft powerplant.

The invention discloses a solid-state high-power switch suitable for a DC distribution looped network of a more-electric aircraft. The solid-state high-power switch comprises a main switch circuit, an auxiliary power supply circuit, an energy absorption circuit, a detection circuit and a single-chip microcomputer control circuit, the main switch circuit is used for supporting bidirectional flow of high-power current; the auxiliary power supply module comprises a switching power supply which is used for outputting the direct-current looped network voltage to direct-current power supply voltage required by each module, and the switching power supply comprises transformer isolation; the detection circuit comprises a current detection unit and a voltage detection unit and is used for detecting current and voltage signals of the circuit; the single-chip microcomputer control circuit is used for receiving measurement data of the detection circuit, carrying out initial processing on the measurement data and driving an IGBT to be switched on or switched off. According to the solid-state high-power switch, the problems that a traditional mechanical switch is low in response speed, prone to abrasion, complex in maintenance and the like are solved through the solid-state switch technology, and the reliability and the operation efficiency of a power distribution system are improved.

The invention relates to the field of heat dissipation of electric aircrafts, and discloses an electric aircraft heat dissipation optimization method and system based on a numerical simulation model. Respectively analyzing the heat dissipation performance and the heat dissipation efficiency of the model simulation electric aircraft in the wind tunnel test and the numerical simulation so as to evaluate the heat dissipation capability change of the electric aircraft in the wind tunnel test and the numerical simulation, and performing comparative analysis based on the heat dissipation performance and the heat dissipation efficiency to obtain a heat dissipation difference coefficient; the heat dissipation difference coefficient in a period of time is analyzed, the heat dissipation control coefficient is calculated, the heat dissipation control threshold value is set, and the heat dissipation control coefficient is compared with the heat dissipation control threshold value, so that the heat dissipation state is continuously kept within the optimal heat dissipation range, comprehensive evaluation of the heat dissipation performance of the electric aircraft is achieved, and the heat dissipation performance of the electric aircraft is improved. The method is beneficial for realizing fine control of the heat dissipation process, and improves the accuracy of heat dissipation optimization control.

An assembly for an aircraft propulsion system includes a propulsor, an engine, and electrical distribution system, and a controller. The propulsor is configured for rotation about a rotational axis. The engine includes a rotor coupled with the propulsor. The electrical distribution system includes an electric motor. The electric motor is coupled with the propulsor. The electric motor and the rotor are configured to cooperatively control rotation of the propulsor about the rotational axis by applying a total torque to the propulsor. The total torque includes a motor torque of the electric motor and an engine torque of the rotor. The controller is configured to: identify a target rotation speed for the propulsor, identify a deviation of an actual rotation speed of the propulsor from the identified target rotation speed, change a target total torque for the propulsor, control the engine to change an actual engine torque of the rotor to the target total torque, and while controlling the engine to change the actual engine torque of the rotor to the target total torque, identify a torque difference between the actual engine torque and the target total torque and control the electric motor to apply a target motor torque to the propulsor based on the torque difference.

In some embodiments, an electric aircraft may include a plurality of flight components including a plurality of control surfaces; a plurality of lift propulsors; at least a thrust propulsor; and a plurality of electric motors configured to power the plurality of propulsors; a flight controller, communicatively connected to a pilot input and flight components. In some embodiments, a flight controller may be configured to receive control datum from a pilot input; and generate an output datum as a function of the control datum.

The embodiment of the invention discloses a GBIC-based multi-electric aircraft starting power generation system DBN model lightweight method, and relates to the technical field of reliability design and modeling of an equipment complex system.The GBIC-based multi-electric aircraft starting power generation system DBN model lightweight method comprises the steps that a DBN model of a starting power generation system is established, nodes in the DBN model correspond to key components and working states of the starting power generation system, and the nodes in the DBN model correspond to the key components and the working states of the starting power generation system; directed edges in the DBN model are used for describing the mutual relation between the nodes, a redundant structure in the DBN model is recognized, lightweight processing is carried out, and the fault state of the starting power generation system is recognized through the DBN model subjected to the lightweight processing. Aiming at the problems of complex reasoning calculation and low efficiency in the use process of the dynamic Bayesian analysis method of the equipment complex polymorphic system, the grey system theory and the Bayesian information criterion are introduced, the lightweight of the dynamic Bayesian network of the complex polymorphic system is realized, the redundant structure is eliminated, and the accuracy of the result is ensured.

An emergency backup power source for an electrically controlled and motorized window shade unit includes an input power switch, and charge and discharge control circuit, and a supercapacitor bank. During normal operation a nominal voltage level from a main power source of the aircraft is provided directly to the window shade unit through the input power switch, and the supercapacitor bank is charged. The input voltage is monitored by the emergency backup power source, which controls the input power switch. When the input voltage drops below a threshold level, the input power switch is opened and the charge and discharge circuit operates to provide an emergency output voltage to the window shade unit from the supercapacitor bank so that the window shade can be fully opened or closed, as is required.

The invention discloses a 12000-turn range extender generator for an electric aircraft, which adopts a mode of 24-slot and 20-pole inner rotor and radial insertion of seven-section rotor iron core skewed slot magnetic steel, and has the excellent performances of symmetrical magnetomotive force waveform, high winding coefficient, low torque ripple and cogging torque and no end ring current. The motor adopts a centralized winding, winding oil cooling and shell water cooling mode, and is suitable for a generator for a range extender of an electric aircraft.

The invention discloses a range-extended electric aircraft power distribution system and a construction method, and the system constructs a graded power distribution network comprising four stages of protection devices through distributing different windings of a multi-winding motor to different load modules and combining the design of dual power sources of a battery pack and a range extender; by means of the design, independent power supply and fault isolation of winding power are achieved, and it is ensured that power is not completely lost under a single fault; and the dual power sources are backups for each other through an interconnection line, so that the power supply reliability is improved, and the power distribution at different flight stages is optimized. By adopting the technical means of multi-winding motor design and four-stage grading circuit protection, the technical problem that the safety and reliability of a power system of an electric aircraft are insufficient due to a single power supply or a motor fault is solved.