317 results about "Electric aircraft" patented technology

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 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.

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.

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 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 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.

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.

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.

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

This disclosure relates to flight control of electric aircraft and other vehicles. A computer-implemented method for estimating available range of an aircraft in flight is disclosed, comprising: receiving electrical information of one or more batteries measured using a first sensor; estimating aircraft-level energy based on electrical information of the one or more batteries; receiving one or more of an altitude of the aircraft or a current airspeed of the aircraft measured using a second sensor; estimating a steady-state force based on the one or more of the altitude of the aircraft or the current airspeed of the aircraft; estimating one or more of a vertical landing range or a horizontal landing range based on the one or more of the estimated aircraft-level energy or the estimated steady-state force; and displaying the one or more of the estimated vertical landing range or the estimated horizontal landing range on a display.

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.

This application provides a method and apparatus for calibrating the rotor pitch of an electric aircraft, belonging to the field of rotor pitch calibration. The method includes: controlling a servo motor to drive the rotor rotation; recording the servo motor travel value during the rotation process; measuring the collective pitch, longitudinal periodic pitch, and lateral periodic pitch of the rotor when the servo motor reaches each travel value; calculating the forward transformation matrix between the servo motor travel value and the three corresponding pitch coefficients using a cubic fitting method; generating training samples using the forward transformation matrix; inputting the training samples into a neural network for training to obtain the inverse transformation relationship between the servo motor travel value and the three pitch coefficients; and calibrating the rotor pitch of the electric aircraft using the aforementioned forward transformation matrix and inverse transformation relationship.

The application discloses a kind of determination methods of maximum take-off weight and battery weight of electric aircraft, wherein the maximum take-off weight determination method of electric aircraft includes: establishing the maximum take-off weight equation of electric aircraft;Solving the maximum take-off weight equation of electric aircraft, obtain the maximum take-off weight of electric aircraft.The determination method of the battery weight of electric aircraft includes: determining the maximum take-off weight of electric aircraft;Maximum take-off weight is used to calculate the battery weight of electric aircraft.The determination method of the maximum take-off weight and battery weight of electric aircraft, the energy consumed by aircraft mission profile is used to estimate the battery weight, the estimation result is more accurate, the reliability is higher, the calculation process is simpler, and the calculation parameter is more simplified.

Electric aircraft charging system including charging cable configured to carry electricity and energy source, wherein the energy source is electrically connected to the charging cable. The system also including a cable reel module, the cable reel module including a reel, wherein the reel is rotatably mounted to the cable reel module, wherein the reel is configured to rotate in a forward direction and a reverse direction, and the charging cable, in a stowed configuration, is wound around the reel. The cable reel module further including a rotation mechanism configured to rotate the reel in reverse, and a cable reel module door having a closed position and an open position, wherein the closed position prevents access to the reel and the open position allows access to the reel. The system additionally including a controller communicatively connected to the rotation mechanism and configured to send a retraction signal to the rotation mechanism.

The invention discloses a modular self-adaptive vibration reduction separation structure and an electric aircraft, and belongs to the technical field of electric aircrafts, and the modular self-adaptive vibration reduction separation structure comprises an electric aircraft connecting plate assembly used for being connected with an electric aircraft body; the electric buffering assembly is arranged below the electric aircraft connecting plate assembly, and one end of the electric buffering assembly is movably connected with the electric aircraft connecting plate assembly; the driving assembly is arranged below the electric aircraft connecting plate assembly, and the driving assembly is connected with one end of the electric buffering assembly; and the separating assembly is arranged below the electric aircraft connecting plate assembly, the separating assembly and the electric aircraft connecting plate assembly are arranged in a spaced mode, the separating assembly is movably connected with the electric buffering assembly, and the separating assembly is connected with the execution end of the driving assembly. According to the invention, the problems of overlarge impact load, poor separation adaptability and insufficient control precision caused by fixed buffer performance of a traditional device are solved.

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

The utility model provides an electric aircraft cold and hot mixing device which comprises a cold air collecting assembly, a mixing assembly and a heat dissipation assembly, the mixing assembly is arranged above a battery assembly, the mixing assembly is provided with a mixing cavity, and an outlet section of the mixing cavity is communicated with a cabin environment control system through a ventilation pipeline; the cold air collecting assembly comprises a plurality of cold air pipelines, cold air inlets of the cold air pipelines are formed in wings or fuselage of the electric aircraft and correspond to the battery assembly, and cold air outlets of the cold air pipelines extend into the inlet section of the mixing cavity; the heat dissipation assembly comprises a plurality of heat conduction coil pipes, the heat conduction coil pipes are arranged on the bottom surface of the battery assembly, hot air inlets of the heat conduction coil pipes are communicated with the battery assembly, and hot air outlets of the heat conduction coil pipes extend into the inlet section of the mixing cavity and are staggered with the cold air pipeline. Natural high-altitude cold air is fully utilized, and the cold air and hot air generated by the battery are fully mixed in the mixing cavity, so that the temperature of the air is effectively reduced, and the heat dissipation efficiency of the battery system is improved.

A system for fault detection and control in an electric aircraft including an inertial measurement unit, the inertial measurement unit including at least a sensor configured to detect a torque datum associated with at least a propulsor. The system includes an observer, the observer configured to generate a torque prediction datum associated with the at least a propulsor, compare the torque prediction datum with the torque datum, and generate a residual datum as a function of the comparison. The system includes a mixer, the mixer comprising circuitry configured to generate, as a function of the residual datum, a torque priority command datum and transmit, to the at least a propulsor, the torque priority command datum.

The application provides a permanent magnet synchronous motor weak magnetic transition area smooth transition method and system based on a reference voltage adaptive reduction, and belongs to the field of permanent magnet synchronous motor control. The method solves the problem of torque fluctuation caused by voltage overshoot in the existing voltage closed-loop weak magnetic control with a fixed reference voltage. The method comprises the following steps: obtaining a theoretical maximum torque according to a preset current limit value by simultaneously solving a torque equation and a current constraint equation; calculating torque ratio parameters and speed ratio parameters in real time; calculating a reference voltage adaptive coefficient based on an adaptive smooth transition algorithm; multiplying the adaptive coefficient by the inscribed circle voltage of an inverter to dynamically adjust the reference value of the voltage closed loop, so that smooth control of the weak magnetic transition area is realized. The method is mainly used in the fields of electric aircrafts, electric vehicles and robots.

There is provided an induction machine (100) comprising a rotor (120); a stator (140); and a phase-shift oscillator (160). The stator comprises: a first winding (141); and a second winding (142), arranged at a first angle (101) relative to said first winding. The phase-shift oscillator comprises: a transistor (170), the transistor (170) being a high-electron mobility transistor, HEMT; and a phase-shift network (180). The first winding is connected to a first node (181) of the phase-shift network and wherein the second winding is connected to a second node (182) of the phase-shift network, wherein the phase-shift oscillator is configured to provide a first phase electric signal at the first node and a second phase electric signal at the second node, wherein a difference between the first and second phase corresponds to the first angle. There is also provided an electric aircraft propulsion system comprising the induction machine.

The invention discloses a hydrogen-lithium hybrid aircraft weight estimation method, and relates to the field of overall design of hydrogen energy aircrafts and hybrid aircrafts, and the method comprises the following steps: carrying out constraint analysis according to aircraft task requirements and existing database information, and obtaining a power-to-weight ratio and a wing load; performing performance analysis on the basis of the power-to-weight ratio, the wing load, the aircraft weight initial value and the flight mission profile to obtain an aircraft power demand and an energy demand; calculating the weight of each part of the airplane based on the airplane power demand, the energy demand, the power mixing degree and the energy mixing degree; and obtaining the total weight of the aircraft based on the weight of each part of the aircraft, judging whether the total weight of the aircraft converges or not, if so, outputting the total weight of the aircraft, and completing the weight estimation of the hydrogen-lithium hybrid aircraft. The method solves the problem that an existing method lacks aircraft weight research on different proportions under the condition that energy is supplied to the lithium hydrogen battery at the same time.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircraft driven by an electric propulsion system, as well as other types of aircraft. In some embodiments, a flight control system of an aircraft is disclosed, the flight control system configured to: determine one or more desired commands for an electric aircraft; determining at least one reference command based on the desired command and one or more aircraft conditions; monitoring an energy state of a plurality of battery packs, wherein at least a first battery pack of the plurality of battery packs is electrically isolated from at least a second battery pack of the plurality of battery packs; adjusting the at least one reference command based on the monitored energy states of the plurality of battery packs; generating a control command for the plurality of actuators based on the adjusted at least one reference command; and controlling the plurality of actuators to satisfy the one or more desired commands of the electric aircraft according to the generated control commands.

In a first aspect, described herein is a direct drive electric aircraft propulsion wherein the propulsion rotor torque is decoupled from the primary proprotor mast moment forces. A hub shaft locates the propulsion proprotor assembly in space relative to the aircraft nacelle, while a motor torque coupler transfers torque from the electric motor to the propulsion rotor while resolving a negligible amount of mast moment through the electric motor.

1. The name of the design product: composite wing electric aircraft. 2. The use of the design product: for manned flight, sightseeing, transportation of goods, rescue, etc. 3. The design points of the design product: in shape. 4. The picture or photo that best indicates the design points: perspective view.

The utility model relates to the technical field of aircraft assembly tools, in particular to an inner and outer double-skin fuselage assembly tool of an electric aircraft, which comprises a fuselage belly tool, an inner and outer double-skin fuselage assembly tool and an outer and inner double-skin fuselage assembly tool, and is characterized in that the fuselage belly tool is used for supporting and fixing a fuselage belly; the fuselage tail tool is used for supporting and fixing the tail of the fuselage; the two skin tools are arranged, the two skin tools are used for supporting and fixing a fuselage left skin and a fuselage right skin correspondingly, and the two skin tools are symmetrically arranged relative to the fuselage abdomen tool so that the two skin tools can be distributed in a separated mode; the two skin tools are symmetrically arranged relative to the fuselage belly tool, so that the two skin tools are distributed in a separated mode, the distance between the two skin tools can be adjusted, and therefore the separated double-skin tool can adapt to fuselage appearance faces of aircrafts of various models.

The invention discloses a permanent magnet synchronous motor sensorless current reconstruction method suitable for an electric aircraft, which comprises the following steps: switching modes of a motor control system of the electric aircraft through a fault detection module and a switch, the modes comprising normal operation and fault tolerance; whether a current sensor fails or not is detected in real time through a fault detection module, if the current sensor fails, a direct-current bus current signal is obtained through single-resistor sampling, another phase of current and rotor position and rotating speed information are reconstructed in real time through an EKF method, third phase of current is calculated in combination with the Kirchhoff's current law, and the third phase of current is calculated. And position-sensorless vector control is realized. According to the invention, by combining the single-resistor sampling technology and the EKF algorithm, the three-phase current, the rotor position and the rotating speed of the motor can be reconstructed at the same time only through a single sampling resistor, the feedback current of the motor can be normally obtained under the condition that the current sensor fails, and the safe flight of an airplane is ensured.