Electric aircraft power plants and associated methods are provided. One power plant includes an emergency power unit (EPU) for providing electric power in the event of a malfunction of a battery pack of an electric aircraft to permit the electric aircraft to make an emergency maneuver. The EPU includes a rocket engine for generating a stream of exhaust fluid using a rocket propellant, a turbine operatively connected to extract energy from the stream of exhaust fluid generated by the rocket engine, and an electric generator operatively connected to be driven by the turbine and to supply electric power to an electric motor propelling the electric aircraft.

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 power distribution controller for a hybrid aircraft is configured to continuously: obtain a state-of-charge (SoC) measurement for a battery of the hybrid aircraft; obtain a fuel level measurement for a secondary energy source of the hybrid aircraft; receive a control input indicating one of a throttle level or an operating mode for one or more motors of the hybrid aircraft; calculate a ratio of energy to source from each of the battery and the secondary energy source in order to operate the one or more motors of the hybrid aircraft based on the control input, the SoC measurement, and the fuel level measurement; and transmit a control signal that causes energy to be apportioned from the battery and the secondary energy source to the one or more motors based on the determined ratio.

The disclosure describes a system that includes an auxiliary power unit (APU), an APU throttle valve, and an environmental control system (ECS) bypass valve. The APU is configured to receive cabin discharge air from an aircraft cabin and receive ECS supply air from an air pressurization system (APS). The APU throttle valve is configured to control flow of cabin discharge air from the cabin to the APU. The ECS bypass valve configured to control flow of ECS supply air from the APS to the APU.

An auxiliary power unit for an aircraft, having a compressor, an intercooler including first conduit(s) having an inlet in fluid communication with the compressor outlet and second conduit(s) configured for circulation of a coolant therethrough, an engine core having an inlet in fluid communication with an outlet of the first conduit(s), and a bleed conduit in fluid communication with the outlet of the first conduit(s) through a bleed air valve. The auxiliary power unit may include a generator in driving engagement with the shaft of the engine core to provide electrical power for the aircraft. A method of providing compressed air and electrical power to an aircraft is also discussed.

A quick attach bracket includes a first portion and a second portion aligned substantially perpendicular with the first portion, a retaining ridge disposed across the first portion and extending in the direction of the second portion, and a hub disposed on the second portion and extending in the direction of the first portion. The retaining ridge is configured to receive a flange of a structure. The hub includes a sidewall that extends from the second portion in the direction of the first portion, a façade aligned with the second portion, and a rounded portion connecting the sidewall and the façade to assist with installation of the hub into a recess or through-hole of the structure. The quick attach bracket is installable by hand via hooking of the retaining ridge on the flange and insertion of the hub into the recess or through-hole.

A boundary layer ingestion-open rotor system for use with an aircraft having a fuselage, wings, and an empennage includes an open rotor assembly, one or more energy storage systems, and an electronic control unit (ECU). The open rotor assembly includes fan blades connected to and extending radially from a rotor hub, and a linkage assembly connecting the hub to the fuselage aft of the empennage within a predefined boundary layer of airflow around the fuselage. The energy storage systems are connectable to the rotor hub. In response to an electronic control signal, the system(s) selectively energize the open rotor assembly to cause rotation of the hub to occur within the boundary layer. The ECU selectively generates the electronic control signals to energize the open rotor assembly during one or more predetermined flight operating phases of the aircraft, e.g., cruise, takeoff, landing, and descent.

An aircraft (41) including a fuselage (14) having a skin (15) and extending along a longitudinal axis (39) from a front end to a rear end (17) of the aircraft, and an engine (30) including an air intake (31) forming an opening in the skin (15) and an exhaust (32) forming another opening in the skin (15). The engine is in a rear fuselage section, in which the exhaust (32) is situated ahead of the air intake (31) along the longitudinal direction (39).

A hybrid propulsion system for an electric aircraft, the system including an electric aircraft including a fuselage. The fuselage including an energy source containing electric power. The electric aircraft further including at least a laterally extending element attached to the fuselage and extending laterally from the fuselage. The electric aircraft further including at least a propulsor electrically connected to the energy source. The system also including at least a power unit pod attached to the at least a laterally extending element and including an auxiliary power unit configured to generate electric power. The power unit pod also including a fuel tank in fluid communication with the auxiliary power unit and a power output line electrically connected to the energy source of the electric aircraft.

The disclosure addresses an external lighting unit for an aircraft. This external lighting unit includes a lighting unit housing have a mounting flange that is configured to provide a reduced contact with a supporting structure (e.g., a tail cone of an aircraft), which in turn reduces conductive heat transfer between the supporting structure and the lighting unit housing (more specifically reduces conductive heat transfer between the supporting structure and a printed circuit board that is disposed within the lighting unit housing and that may be seated on a portion of the lighting unit housing). Open spaces on the surface of the mounting flange that face toward the supporting structure may be occupied by a thermal insulator.