447 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.

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.

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 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.

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.

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.

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 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.

A system for flight control configured for use in an electric aircraft includes an inertial measurement unit (IMU) and configured to detect an aircraft angle and an aircraft angle rate. The system includes a flight controller including an outer loop controller configured to receive the input datum from the sensor, receive the aircraft angle from the IMU, and generate a rate setpoint as a function of the input datum. The system includes an inner loop controller configured to receive the aircraft angle rate, receive the rate setpoint from the outer loop controller, and generate a moment datum as a function of the rate setpoint. The system includes a mixer configured to receive the moment datum, perform a torque allocation as a function of the moment datum, and generate a motor command datum as a function of the torque allocation.

The utility model relates to the technical field of electric aircrafts, in particular to a heat management system of an electric aircraft, which comprises a heat pump unit and a heat pipe radiator, and the heat pump unit comprises a compressor, a one-way valve, a pipeline, a heat management controller and a valve; the heat pipe radiator comprises a copper heat pipe and aluminum fins. The utility model aims to realize effective heat dissipation of heat generated by the electric aircraft by optimizing a heat management system, and meanwhile, the heat is used for heating the interior of an aircraft cabin, so that the comfort of passengers is improved. In addition, the system further has an intelligent heat management function, and it is ensured that the electric aircraft can keep the optimal heat balance state under various flight conditions.

A thermal management system for an aircraft engine includes a closed-loop fluid circuit including a coolant supply leg, a first electric power system leg, and a second electric power system leg. A coolant supply system is fluidly coupled to the coolant supply leg for providing a coolant fluid to the coolant supply leg, the first electric power system leg, and the second electric power system leg of the closed-loop fluid circuit. A flow control device fluidly couples the coolant supply system via the coolant supply leg, to the first electric power system leg, and to the second electric power system leg. A first heat source is in thermal communication with the coolant fluid via the first electric power system leg, and a second heat source is in thermal communication with the coolant fluid via the second electric power system leg.

The present disclosure is generally directed to a flight control system and method of flight control for an electric aircraft. The system includes a pilot input communicatively connected to an electric aircraft, wherein the pilot input is configured to receive an input datum, and plurality of flight components communicatively connected to the electric aircraft, wherein the plurality of flight components includes a plurality of control surfaces. The system also includes a flight controller, wherein the flight controller is configured to determine a phase of flight, determine a command datum to control a position of the plurality of control surfaces as a function of the input datum, and command, when the phase of flight is determined to be hover, the plurality of control surfaces using the command datum.

An electrically operated aircraft deicing vehicle includes a chassis having longitudinally extending beams. The chassis supports a rechargeable battery system including one or more battery modules and a charging unit, front and rear axles, a positioning system for a fluid spraying system and a fluid tank system containing deicing fluid. An electric drive motor is operably coupled to the battery system. The electric drive motor is positioned aft of the front axle and between the longitudinally extending beams of the chassis. The electric drive motor is operable to rotate the front wheels. A hydraulic pump system is powered by the battery system. The hydraulic pump system is operably coupled to the positioning system enabling movement of the fluid spraying system. A heat transfer unit is positioned forward of and is operably coupled to the battery system for thermally conditioning the battery system.

The invention discloses an electric aircraft high-voltage power distribution control method and system, a medium and a product thereof, and relates to the technical field of high-voltage power distribution, and the method comprises the steps: receiving a DC power supply outputted by a power battery pack, and achieving the power-on soft start through a pre-charging loop; distributing electric energy to a motor controller, an auxiliary power supply module, an air conditioning system and an electric heating system through a power distribution level according to the load priority; the power distribution stage integrates multiple solid-state relays or contactors to form load branches. An independent protection unit is configured for each load branch at a load output stage, and the independent protection units monitor voltage, current and temperature parameters in real time and execute fault isolation operation when the monitored parameters exceed corresponding preset thresholds. According to the invention, high-voltage power distribution control with load priority scheduling and branch-level rapid fault isolation is realized under the condition of high power density.

In an aspect, a system for propeller parking control for an electric aircraft. The system include at least a sensor and a computing device. A sensor may be configured to generate angular datum. The computing device may be configured to generate a trajectory as a function of angular datum. The computing device may also be configured to initiate the transition from hover to fixed-wing flight as a function of a trajectory.

The invention discloses an electric aircraft cross compass attitude control system, and relates to the technical field of unmanned aerial vehicle attitude control, the electric aircraft cross compass attitude control system comprises # imgabs 0 # rotor motors and further comprises an annular base frame, and # imgabs 1 # rotor motors form an annular array based on the annular base frame; a controller; and the driving mechanism is used for receiving the driving electric signal of the controller and adjusting the position of the rotor motor. According to the cross compass attitude control system of the electric aircraft, when the function of the rotor motor is abnormal or damaged through automatic monitoring, the abnormal or damaged rotor motor is abandoned, and the position or the rotating speed of the rotor motor is correspondingly adjusted based on a corresponding strategy, so that the rotor motor annularly moves relative to the annular base frame; according to the invention, the attitude basis conforming to aerodynamics is reconstructed to optimize lift distribution to compensate imbalance, so that the aircraft can stably fly to the maximum extent in an extreme environment, high controllability is kept, and a better use prospect is brought.

The invention discloses a battery module based on a phase-change material and an air-cooling heat dissipation system and an electric aircraft. The battery module comprises a module shell frame and a heat conducting plate, the module shell frame is formed by fixedly connecting four side walls; the plurality of heat conducting plates are arranged in parallel at intervals between the module shell frames, and battery cell cabins, phase change material cabins and heat dissipation air channels which are alternately distributed are formed between the adjacent heat conducting plates; the battery cell cabin is used for accommodating a battery cell in direct contact with the heat conducting plate; the phase-change material cabin is filled with a phase-change material and is in heat conduction connection with the battery cell through a heat conduction plate; and the heat dissipation air channel is communicated with an external air cooling channel. According to the invention, the phase change material cabins and the air cooling heat dissipation air channels are alternately stacked and arranged, and a three-stage heat conduction path of battery cell-phase change material-air channels is formed through separation, so that the balance of efficient absorption and continuous heat dissipation of short-time high-power heat production is realized; the take-off stage depends on rapid heat absorption of the phase-change material, and the cruise stage takes away heat through high-speed airflow and promotes solidification and regeneration of the phase-change material, so that the problem of continuous energy consumption of the system is solved.

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 small sample electric aircraft composite material combined structure dynamics modeling method based on attention element learning. Comprising the following steps: step 1, a used data set comes from vibration experiments performed on a composite material laminated wing box structure, a composite material laminated air beam structure and a composite material conical-cylindrical combined shell under different temperature conditions; under the condition of limited data, various composite material structures and environment temperatures are divided into independent task samples, each task comprises vibration test data of the structure at the respective temperature, and the data are divided into a meta training set and a meta test set; the invention relates to the technical field of dynamic modeling, meta-learning and dynamic modeling of an electric aircraft composite material combined structure are creatively fused together, the AMLT model can utilize the rapid task adaptability and the multi-task learning ability of MAML to rapidly transfer and apply knowledge to a new task, and the modeling efficiency is improved. And the method has higher adaptability and generalization ability when facing new tasks.