A method to modify an aircraft including: disconnecting a first engine control device from command cables and transmission cables, wherein the first engine control device is in a fuselage section forward of a pressure bulkhead; replacing the engine control device with a jumper connector positioned in the fuselage section, wherein the jumper connector electrically connects the command cables to the transmission cables; installing a second engine control device in the fuselage aft of the pressure bulkhead, wherein the second engine control device is in an engine compartment of the fuselage; connecting the second engine control device to transmission cables at a location at or near the pressure bulkhead; connecting sensor cabling directly to the second engine control device, and connecting the engine control device directly to the engine.

A method and power distribution system for operating in a low power consumption mode includes a primary power distribution node defining a primary distribution switch having an output and operable in a first conducting mode and a second non-conducting mode, and wherein operating in the second non-conducting mode includes a leakage current through the power distribution switch, at least one enabled electrical load downstream of the primary power distribution node, the at least one enabled electrical load connectable to the primary power distribution node by way of the primary distribution switch, and a primary power distribution node power source configured to supply power to the output of the primary distribution switch when the primary distribution switch is operating in the second non-conducting mode.

An aircraft turbine engine includes a gas generator having a longitudinal axis (A), a fan located at an upstream end of the gas generator and configured to turn about said axis, and an electric machine of generally annular shape. The electric machine is coaxially mounted downstream of the fan and has a rotor coupled to rotate with the fan. The electric machine further includes a stator connected to an electronic power circuit by at least one rigid electrically-conductive bar.

The invention relates to an electric architecture for a twin-engined, hybrid thermal/electric propulsion aircraft and, for each turboshaft engine, the architecture comprises: —a high-voltage DC propulsive electric distribution network (32), —a non-propulsive electric distribution network (56) which is connected to loads of the aircraft, and—an electric distribution network (76) which is connected to loads of an electrified control system of the turboshaft engine, and wherein power supply sources are shared for these different networks.

A propulsion system comprising a nacelle with an air channel along a longitudinal direction, an electric motor whose output drives a propeller, and a fuel cell, comprising a core outside the air channel, open channels, each of which has an inlet and an outlet opening in the air channel, and, for each open channel, a fuel chamber, an electrolyte between the open channel and the fuel chamber, a cathode, and an anode, each open channel having an inlet surface area which is less than the surface area of an intermediate area between the inlet and the outlet, the surface area of the outlet being smaller than the surface area of the intermediate area. Such a system makes it possible to have the fuel cell close to the electric motor, thereby reducing the lengths of the electrical conductors between them, and consequently improving the operation of the fuel cell.

A device for electrically connecting cables where a change of direction occurs. The device includes at least two busbars as well as a plate having independent internal ducts in which the busbars are located, each duct being electrically insulated from the other ducts, the busbars ensuring electrical conduction from one open end of a duct to the other end. The cables are fastened at these open ends and thus do not undergo any deformation, the change of direction taking place inside the plate with the aid of the busbars.

A system and method for a recharging station including an elevated landing pad, a rechargeable component coupled to the elevated landing pad, a power delivery unit configured to deliver power from a power supply unit or power storage unit to the recharging component, and a support component coupled to the bottom of the elevated landing pad.

The disclosure relates to a modular arrangement of a converter for the electrical supply of a multi-phase electric motor. The modular arrangement includes multiple output stage modules, each module having a separate first housing. An output stage module is provided for each phase of the electric motor, wherein the output stage module has a phase intermediate circuit and an inverter circuit, together forming a commutation cell. The phase intermediate circuit supplies DC voltage and is electrically connected to the inverter circuit. The disclosure also relates to a motor arrangement and an aircraft having a motor arrangement of this type.

An electric aircraft includes a propeller, an electric motor, and a controller. The electric motor is configured to supply power to the propeller. The controller is configured to control the electric motor. The controller is disposed inside a leading-edge portion of a wing to cause heat to be transmitted to a skin of the wing. The heat is generated by the controller when the controller controls the electric motor.

Methods and apparatus for detecting and characterizing arc faults in an aerospace electric propulsion system and then coordinating the operation of various elements of the protection system to execute a fault-clearing sequence. In a current-based method, the arc is detected and characterized based on differential readouts from current sensors. The difference between currents measured at two ends of a protection zone are compared to a difference threshold. In a power-based method, the arc is detected and characterized based on differential readouts from voltage and current sensors. The differential voltage and current readouts are used to compute the respective powers at two ends of a protection zone. The difference between the respective powers is integrated over a period of time and then the integrated difference is compared to a difference threshold. A differential protection trip mode is invoked when the difference threshold is exceeded.