59 results about "Electric aircraft" patented technology

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 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.

A propulsion system includes a propulsor, an engine, an electrical assembly, a controller, and an uncontrollable high thrust (UHT) protection assembly. The engine includes a fuel system. The engine is coupled with the propulsor. The fuel system includes a fuel control unit. The electrical assembly includes an electric motor and a motor control unit. The electric motor is coupled with the propulsor. The motor control unit is electrically connected with the electric motor. The controller includes a first controller channel. operable to control the fuel control unit to direct fuel to the combustor and to control the motor control unit to direct electrical power to the electric motor. The UHT protection assembly includes a first protection circuit operable to identify a presence or an absence of a UHT condition of the propulsor and, in response, generate a shutdown signal for the fuel control unit and the motor control unit.

The embodiment of the specification provides a kind of based on quasi-static electric quantity discrete electric aircraft reachable domain evaluation method and system, wherein, method includes: the quasi-static dimension reduction conversion of electric aircraft dynamics model, according to quasi-static hypothesis, simplify electric aircraft dynamics equation, it is decoupled as horizontal plane kinematics equation and push resistance balance and lift weight balance equation constraint, reduce state quantity dimension, obtain dimension reduction flight dynamics model;The electric quantity discrete conversion of dimension reduction flight dynamics model, the form of relative to time dynamics equation is converted into relative to electric quantity consumption value, determines simulation interval and discrete grid;Based on the dimension reduction flight dynamics model, by setting roll angle numerical sequence, numerical integration simulation is carried out on the simulation interval and discrete grid, and the horizontal plane maneuvering reachable domain boundary of electric aircraft is obtained.

1. Name of the designed product: electric airplane. 2. Use of the designed product: for airplane. 3. Design points of the designed product: in shape. 4. Picture or photo that best shows the design points: perspective view 1.

This application provides a method and system for determining risk sources of distributed electric aircraft based on flight profiles, applied in the field of aircraft simulation technology. It includes: acquiring the overall construction content of the flight profile and topology of the distributed electric aircraft, and obtaining typical operating conditions for each stage in multiple phases corresponding to the flight profile. The overall construction content includes a high-power-density battery pack, a DC-DC buck converter, an electronic speed controller, a DC brushless motor, and rotors / propellers; simulating the mathematical model of the overall construction content under the target typical operating conditions of the target phase, obtaining simulation results for the target typical operating conditions; and accurately determining the flight test risk sources of the distributed electric aircraft under the target typical operating conditions from the high-power-density battery pack, DC-DC buck converter, electronic speed controller, DC brushless motor, and rotors / propellers based on the simulation results. This significantly reduces the flight risk of new aircraft configurations (i.e., distributed electric aircraft) and improves flight test efficiency.

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.

This application relates to a ducted-coupled propulsion system for an electric aircraft and an electric aircraft. The ducted-coupled propulsion system includes a casing, fan blades, a shaft coupling assembly, and an electric drive assembly. The fan blades are installed inside the casing, and the shaft coupling assembly is connected to the fan blades. The electric drive assembly includes a first motor and a second motor, which are independently configured. The first motor includes a first motor housing and a first output shaft, and the second motor includes a second motor housing and a second output shaft. Both ends of the first and second motor housings are open along their axial directions, so that their internal cavities are connected to the duct. The first and second output shafts are used for transmission connection with the shaft coupling assembly, so that the power generated by one or both can be transmitted to the fan blades. This ducted-coupled propulsion system has power redundancy backup and peak thrust superposition capabilities, and utilizes duct airflow to achieve self-pumped air cooling, eliminating the need for additional cooling fans or liquid cooling systems, thus achieving the effects of weight reduction, structural simplification, and improved reliability.

Embodiments provide an electric aircraft with a plurality of lift fan assemblies that are configured to provide vertical lift, and one or more pusher propellers that are configured to provide forward thrust. The lift fan assemblies may be coupled to the wings of the aircraft via one or more support structures, and the wings may be coupled to an upper region of the fuselage. The pusher propeller(s) may be coupled to a tailing end of the fuselage. The lift fan assemblies and the pusher propeller(s) may provide thrust and movement in directions that are orthogonal to one another. A control system coupled to the aircraft may control the lift fan assemblies and the one or more pusher propellers to activate, increase in power, and decrease in power. The lift fan assemblies and the one or more pusher propellers may be operated separately, and may be active at different times.

To provide a mobile component that enhances durability while suppressing the increase in weight of the electric motor. [Solution] According to one aspect of the present invention, a moving element for an electric motor that moves in a specific direction of movement is provided, comprising a first magnet and a second magnet arranged adjacent to the first magnet in the direction of movement, wherein the first magnet and the second magnet are in contact with each other, thereby suppressing the movement of the first magnet and the second magnet in a direction intersecting the direction of movement.

The present disclosure relates to power and onboard electrical power systems for all-electric aircraft and related power supply systems. A power electrical power system includes a plurality of batteries for providing electrical power, a plurality of DC / DC converters, each of which is electrically connected to a respective one of the plurality of batteries to DC / DC convert direct current output by the battery, a plurality of distribution panels, each of which is electrically connected to an output of at least one of the plurality of DC / DC converters to receive DC / DC converted direct current, and a plurality of DC / AC inverters, each of which is electrically connected to an output of one of the plurality of distribution panels to convert direct current output from the distribution panel to three-phase alternating current, thereby providing electrical power to a drive device of the all-electric aircraft.

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.

An electric aircraft comprising: a fuselage; one or more wings mounted on the fuselage; a plurality of propellers pivotable between lift and forward propulsion configurations; and a plurality of battery packs, wherein at least two of the plurality of battery packs are configured to power: at least a portion of a first propeller of the plurality of propellers and at least a portion of a second propeller of the plurality of propellers, which is different from the first propeller.

This invention discloses a method and system for estimating the State of Charge (SOC) of an electric aircraft battery. The method first constructs a second-order RC equivalent circuit model and sets up a two-dimensional lookup table module. By inputting the real-time operating temperature and the estimated SOC value into the lookup table module, the corresponding open-circuit voltage value is obtained and assigned to the equivalent circuit model. Then, based on this model, an extended Kalman filter algorithm is applied to estimate the SOC value in real time. During algorithm execution, the real-time current and temperature of the battery are monitored, and the process noise covariance matrix and observation noise covariance matrix are dynamically adjusted according to the current change rate and temperature change rate to optimize the SOC estimation result. This invention effectively overcomes the technical defects of existing second-order RC models, such as significant polarization effects and insufficient dynamic response under extreme temperatures, through a temperature-compensated two-dimensional lookup table module. Combined with an adaptively adjusted extended Kalman filter algorithm, it significantly improves the estimation accuracy and reliability of the electric aircraft battery management system under complex operating conditions.

This invention discloses a motor redundancy control system suitable for manned electric helicopters. The system includes: a current sensor, a three-phase voltage sensor, a motor position sensor, an inverse converter module, a motor control algorithm module, a current fault diagnosis and reconstruction module, a Clarke voltage inverter module, a speed and position monitoring module, and a position / speed signal fault diagnosis and reconstruction module. Applying this system ensures the flight safety of manned electric helicopters, preventing crashes due to power loss. This control system primarily addresses the current and position sensors, which are prone to failure during cruise in manned electric helicopters. Different redundancy control methods are proposed for each. When current and position sensors malfunction, the system can quickly diagnose the fault and reconstruct the position and current information using position and current reconstruction methods respectively. The fault information is then uploaded to the flight assistance display and control system to alert the pilot, allowing for an emergency landing at a suitable location, thus ensuring the stable operation of the manned electric helicopter's motor control system.

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

A system and method for a battery power management system for an electric aircraft are disclosed. The system includes at least a flight component of an electric aircraft, at least a battery, at least the battery configured to supply power to at least the flight component of the electric aircraft, at least a sensor communicatively connected to the battery, and at least a controller communicatively connected to the sensor. The controller is configured to receive sensor data from at least the sensor, identify the battery state as a function of the sensor data and a battery threshold, control power from at least the battery to at least the flight component of the electric aircraft as a function of the battery state, and further include reducing torque to at least the flight component.

1. The name of the design product: composite wing electric aircraft (2). 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 1.

This invention discloses a 30,000 rpm range extender generator for electric aircraft. It adopts an internal rotor with 18 slots, 10 poles, and a five-segment rotor core with radially inserted skewed magnets. It has excellent performance characteristics such as symmetrical magnetomotive force waveform, high winding coefficient, low torque pulsation and cogging torque, and no end circulating current. The motor adopts a centralized winding, winding oil cooling, and shell water cooling method, and is suitable for use as a range extender generator for electric aircraft.

Systems and techniques for state control of a flight component of an electric aircraft that includes a high voltage connection to a power supply of the electric aircraft and a communication connection to a communication system of the electric aircraft. The flight component is configured with a hardware circuit designed and configured to receive and / or read a signal on the communication system, compare an identifier of the signal against a stored identifier of the flight component, and in response to the identifier matching the stored identifier, cause the flight component to change from a first state to a second state.

The application provides an electric aircraft thermal runaway emergency response method, system, medium and electric aircraft, and relates to the technical field of aircraft emergency response. The method comprises the following steps: in response to detecting that the electric aircraft has a thermal runaway fault, triggering a thermal runaway alarm; in response to the thermal runaway alarm being triggered, the electric aircraft automatically performs a landing operation to land at a landable location at a predetermined thermal runaway descent rate. The application realizes early warning and accurate response of the fault by constructing a rapid diagnosis and automatic disposal integrated process for battery thermal runaway, can significantly shorten the exposure time of the aircraft in a high-risk state, effectively curb the chain disaster caused by thermal runaway, and thus maximally guarantee the safety of personnel and the aircraft.

This invention discloses a dynamic path planning method for electric aircraft for low-altitude patrol, belonging to the field of low-altitude aircraft planning technology. The method includes constructing a composite environmental situation map containing a terrain obstacle grid map and an airspace access cost map; extracting potential feasible route segments from the current position to the target patrol point and performing energy consumption prediction modeling to obtain the estimated energy consumption value for each segment; comparing this value with the remaining battery capacity percentage; eliminating route segments exceeding the energy budget; and retaining a set of candidate routes that meet the constraints. The candidate routes are then smoothed to eliminate curvature abrupt changes, generating geometrically continuous executable route geometry. A power consumption rate curve is superimposed on the path to calculate the theoretical endurance and assign a time attribute label. A comprehensive score is then combined with the patrol task priority weight to select the optimal dynamic planning path. This method ensures that the path matches the aircraft's energy state, improving the flight adaptability and planning rationality of the route.

An electric aircraft is described, comprising a fuselage and at least one wing attached to the fuselage. The electric aircraft includes a plurality of tilting propellers attached to the at least one wing, the tilting propellers being configured to move between a vertical lift position and a forward flight position. The electric aircraft includes a plurality of tilting mechanisms coupled to at least one tilting propeller. The electric aircraft includes a first actuator coupled to a first subset of the tilting mechanisms. The first subset of tilting mechanisms is identified among the tilting mechanisms according to a coupling scheme. The first actuator simultaneously tilts a first group of tilting propellers coupled to the first subset of the tilting mechanisms.

This invention discloses a wind-powered aircraft and its control method, belonging to the field of new energy aviation technology, and is adaptable to narrow-body and wide-body civil aircraft as well as medium and large-sized UAVs. The invention retains the original fuselage frame and cabin / payload layout of the aircraft, symmetrically arranging compatible horizontal-axis wind turbines at the original engine locations on the wings. Modular power battery energy storage modules store electricity, and multi-motor drive modules provide flight thrust and attitude control power. Medium and large-sized UAVs can be equipped with electric hubs for zero-emission ground taxiing. A distributed flight control module dynamically allocates power according to flight stage and aircraft type, adjusts wind turbine and motor parameters in real time, and sets differentiated fault protection. This invention integrates wind power generation, energy storage, and multi-motor drive, adapting to all types of aircraft by adjusting component specifications. It solves the problems of high carbon emissions from fuel-powered aircraft, short range of electric aircraft, and insufficient integration of wind power with aviation, possessing advantages such as zero carbon emissions, long range, strong compatibility, low cost, and safety and reliability, providing a universal technical solution for new energy aviation.

The present disclosure relates to an aircraft comprising: a fuselage and at least one wing arranged in a tube-and-wing design; a plurality of propulsors; and a primary energy source configured to power the propulsors, the primary energy source comprising rechargeable batteries. The aircraft has a maximum take-off mass ‘MTOM’ of at least 8,618 kg; and an electric range factor ‘ERF’ of at least 6.

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.

An electric aircraft comprising: a fuselage; one or more wings mounted to the fuselage; a plurality of first electric propulsion units arranged on both sides of the fuselage behind the one or more wings during forward flight; a plurality of second electric propulsion units arranged on both sides of the fuselage in front of the one or more wings during forward flight; a plurality of battery packs, each battery pack being configured to power: a partial motor of one of the first electric propulsion units on one side of the fuselage, and a partial motor of one of the second electric propulsion units on the other side of the fuselage.

This invention relates to the field of power control technology for electric aircraft and discloses a power mode switching system for electric aircraft, comprising: a power acquisition module for acquiring full-dimensional state parameters, flight condition parameters, and fault signals from multiple power sources; and a data preprocessing and fusion module for performing noise reduction, calibration, and multi-source fusion analysis on the parameters to construct a unified "power-condition" data model and predict load power changes and the upper limit of power output within the next 1-3 seconds. It possesses advantages such as proactive predictive switching of power modes, optimal energy efficiency under all operating conditions, and rapid fault redundancy replenishment, improving the flight safety, endurance, and power supply reliability of electric flight platforms. It solves the problems of passive switching, poor multi-source coordination, insufficient fault redundancy, and lack of cross-system coordination in existing electric flight power mode switching systems.

This invention specifically relates to a distributed power management strategy for a hybrid electric propulsion system for electric aircraft, belonging to the field of aviation electrification. The system consists of a fuel cell, an energy storage battery, and a supercapacitor connected in parallel to a DC bus via a power converter. In this invention, each unit is independently controlled based on local information, eliminating the need for a communication network: the fuel cell adaptively switches between voltage and power modes according to the load power and battery state of charge, avoiding heavy or light load operation; the battery adjusts its steady-state operating point through a virtual resistance droop mechanism, absorbing steady-state power deficits or excess power; the supercapacitor absorbs dynamic power through a virtual capacitance droop mechanism, suppressing load fluctuations. This invention effectively suppresses the negative impacts of heavy / light loads, load fluctuations, and frequent start-stop cycles on fuel cell lifespan and efficiency, reduces hydrogen consumption, and improves overall system reliability.

The application discloses a kind of aircraft autopilot perception network training data set construction method and system.The method comprises: determining the scene richness requirement of data set collection;Each individual scene requirement and expected data proportion are refined;Each sample is combined category statistics, determines data volume according to existing data set information and project scene requirement, and is sampled based on Monte Carlo method;Actual scene building and data collection are carried out according to the sampling result.The application is based on manned electric aircraft autopilot neural network training data set scene richness requirement, designs a kind of scene richness building method, combines each scene, determines the sample quantity that needs to be collected, finally collects a large amount of sample data, and overcomes the defect that intelligent driving data set is insufficient for low-altitude flight.