A method of in-flight operational assessment for an electric aircraft comprising detecting by a sensor an electrical parameter of an energy source. The method further includes receiving by a controller the electrical parameter from the sensor and determining a power-production capability of the energy source, using the electrical parameter. The method further includes calculating, by the controller, a projected power-consumption need of the electric aircraft and comparing the determined power-production capability of the energy source to the projected power-consumption need. The method includes generating a power production command datum as a function of the comparison of the power-production capability and the projected power-consumption need.

An electric aircraft having fixed pitch lift includes a plurality of flight components, wherein the plurality of flight components further comprises at least a lift propulsor component, wherein the lift propulsor component comprises a plurality of blades configured at an angle of attack, and a flight controller, wherein the flight controller is configured to calculate a flight element using an intermediate representation, and transmit the flight element to the plurality of flight components.

In a control device including a control unit that controls an EPU for powering a rotor included in an electric aircraft, the EPU includes a motor, an inverter, a control circuit, and a first sensor that detects an operation-related value. If there is no a sudden change in the operation-related value detected by the first sensor, the control unit controls the EPU based on the operation-related value detected by the first sensor. When the sudden change occurs, the control unit stops using the operation-related value detected by the first sensor to control the EPU. After the occurrence of the sudden change, when the sudden change ends before a predetermined period elapses, the control unit resumes using the operation-related value detected by the first sensor to control the EPU, and when the sudden change has not ended, the control unit continues the stoppage.

In a control device including a control unit that controls an EPU for powering a rotor included in an electric aircraft, the EPU includes a motor, an inverter, a control circuit, and a first sensor that detects an operation-related value. If there is no a sudden change in the operation-related value detected by the first sensor, the control unit controls the EPU based on the operation-related value detected by the first sensor. When the sudden change occurs, the control unit stops using the operation-related value detected by the first sensor to control the EPU. After the occurrence of the sudden change, when the sudden change ends before a predetermined period elapses, the control unit resumes using the operation-related value detected by the first sensor to control the EPU, and when the sudden change has not ended, the control unit continues the stoppage.

A secondary battery includes an electrode wound body, a first electrode current collector plate, a second electrode current collector plate, an electrolytic solution, and a battery can. The first electrode current collector plate faces a first end face of the electrode wound body. The second electrode current collector plate faces a second end face of the electrode wound body. In the electrode wound body, an outermost wind part of a second electrode is located on an outer side relative to an outermost wind part of a first electrode. A sidewall part of the battery can includes a thin part, and a thick part that protrudes toward an inner side of the battery can along a radial direction. The thick part is located to overlap, in the radial direction, an end part, of the first electrode covered part, located on a side of the first end face in the first direction.

A secondary battery includes an electrode wound body, a first electrode current collector plate, a second electrode current collector plate, an electrolytic solution, and a battery can. The first electrode current collector plate faces a first end face of the electrode wound body. The second electrode current collector plate faces a second end face of the electrode wound body. In the electrode wound body, an outermost wind part of a second electrode is located on an outer side relative to an outermost wind part of a first electrode. A sidewall part of the battery can includes a thin part, and a thick part that protrudes toward an inner side of the battery can along a radial direction. The thick part is located to overlap, in the radial direction, an end part, of the first electrode covered part, located on a side of the first end face in the first direction.

Various implementations directed to an aircraft thermal management system are provided. In one implementation, an aircraft may include a fuselage having one or more fuselage sections. The aircraft may also include one or more electric motors configured to drive one or more propulsion systems of the aircraft, where the one or more electric motors are configured to generate thermal energy. The aircraft may further include an aircraft thermal management system configured to transfer the thermal energy generated by the one or more electric motors to the one or more fuselage sections.

A method and system for providing corrective action to a rotorcraft experiencing motor failure is provided. Included in the method and system are embodiments that receive feedback from sensors directed at measuring a state of motors used to provide lift to the rotorcraft. The method and system also describe embodiments for determining that there is a malfunctioning motor, and furthermore, the appropriate corrective action for responding to the malfunctioning motor. In some embodiments, the method and system are configured to reduce power to the malfunctioning motor while simultaneously adjusting power supplied to the remaining motors such that changes in total thrust and net torque are minimized.

The disclosure generally pertains to a vertical take-off and landing (VTOL) aircraft comprising a fuselage and at least one fixed wing. The aircraft may include at least two powered rotors located generally along a longitudinal axis of the fuselage. The rotor units may be coupled to the fuselage via a rotating chassis, which allows the rotors to provide directed thrust by movement of the rotor units about at least one axis. By moving the rotor units, the aircraft can transition from a hover mode to a transition mode and then to a forward flight mode and back.

As a component of an aerial vehicle, a power management system having an active power control device is provided. The system includes a solar cell converting solar energy into electric energy; a fuel cell provided in the aerial vehicle and converting fuel energy into electric energy by electrochemical reaction; a battery compensating for a lack of electric power supplied from the solar cell and the fuel cell to the aerial vehicle and storing surplus electric power; and an active power control device connecting with all the solar cell, the fuel cell and the battery and combining and distributing electric power generated in the solar cell, the fuel cell and the battery to loads. The system efficiently distributes the power from the respective power sources through the controllable output of the fuel cell in accordance with power required by the aerial vehicle and the solar cell's performance depending on weather.