An aircraft includes a chamber (1), a processor, a memory, and a compressor system (12b) in fluid communication with the chamber. The compressor system (12b) configured to selectively pressurize the chamber (1). The chamber supports a fuel cell (26), a motor, and/or electrical components that electrically communicate with the fuel cell (26) and the motor to power the aircraft. The memory includes instructions stored thereon, which when executed by the processor, cause the aircraft to receive an altitude value of the aircraft, and selectively pressurize the chamber using the compressor system based on the received altitude value to reduce corona discharge in the chamber.

A hybrid electric propulsion (HEP) system of an aircraft includes a gas turbine engine configured to generate rotational power, and an electric propulsion system configured to generate at least one of thrust or lift for operation of the aircraft. The propulsion system includes a propulsor and an electric motor configured to drive the propulsor. A controller is in signal communication with the gas turbine engine and the electric propulsion system. The controller operates the gas turbine engine based on an on-board engine model (OEM), monitors electrical parameters of the electric propulsion system, and updates the OEM in response to changes to the electrical parameters.

A propulsion system for a hybrid electric vehicle comprises a traction motor having first and second stator windings; a power source having a DC power output coupled to the first windings; a battery having a DC power output coupled to the second windings; and a controller to independently control: (i) a first power level output at the first DC power output, and (ii) a second power level of motive power output by the traction motor; wherein responsive to a signal to set the second power level less than full capacity of the traction motor, the controller provides a power difference between the first and second power levels from the second windings to the battery.

An electric propulsion system of an aircraft includes an electrical generator and a cooling device of the electrical generator. It further includes at least one thermoacoustic engine and a heat transfer circuit configured to transport heat dissipated by the electrical generator to the thermoacoustic engine. The cooling device of the electrical generator is at least partially powered by energy from the thermoacoustic engine.

A cooling architecture for an integrated hydrogen-electric engine having a radiator and a hydrogen fuel cell includes a t and a manifold. The turbine is disposed in fluid communication with the hydrogen fuel cell. The turbine is configured to compress a predetermined amount of air and direct a first portion of the predetermined amount of the compressed air to the fuel cell for generating electricity that powers the integrated hydrogen-electric engine. The manifold is disposed in fluid communication with the turbine and positioned to direct a second portion of the predetermined amount of compressed air to the radiator for removing heat from the radiator.

The invention relates to an electric drive for an aircraft hybrid system. This electric drive comprises a rotor and a stator, wherein the stator may be connected to a structure of the aircraft and the rotor has an annular flange with a shaft through opening for mounting on a propeller flange, wherein the flange is formed of at least two parts, wherein each of these parts of the flange delimits a section of the shaft through opening.

A hybrid-electric aerial vehicle is disclosed comprising: an airframe; a plurality of longitudinal booms extending radially from the airframe; a passively charged internal combustion engine operatively coupled with a fuel tank, a generator operatively coupled with the passively charged internal combustion engine; a battery bank operatively coupled with the generator; and a plurality of motors. The passively charged internal combustion engine has an intake engine valve, an exhaust engine valve, and a combustion chamber, wherein the intake engine valve is delayed to provide an expansion ratio in the combustion chamber that is greater than a compression ratio in the combustion chamber. Each of said plurality of motors may be positioned at a distal end of one of said plurality of longitudinal booms and be operatively coupled with a propeller, wherein the plurality of motors is electrically coupled with the battery bank and the generator.

A propulsion system for an aircraft is provided having an aft engine configured to be mounted to the aircraft at an aft end of the aircraft. The aft engine includes a fan rotatable about a central axis of the aft engine having a plurality of fan blades attached to a fan shaft. The aft engine also includes a nacelle encircling the plurality of fan blades and a structural support system for mounting the aft engine to the aircraft. The structural support system extends from the fuselage of the aircraft, through the fan shaft, and to the nacelle when the aft engine is mounted to the aircraft. The aft engine may increase a net thrust of the aircraft when mounted to the aircraft.

The aircraft (10) includes: at least one electric motor (135); at least one stand-alone electrical power supply (110, 120) supplying power to the electric motor; propulsion elements (130) referred to as auxiliary propulsion elements, included in the group including: a stand-alone electrical power supply (130) supplying power to the electric motor, a power supply converting thermal energy into electrical energy and supplying power to the electric motor, and an internal combustion engine; and a structure (100) configured to integrate each electric motor, each stand-alone electrical power supply and the auxiliary propulsion elements, the parameters of the structure being substantially identical regardless of the auxiliary propulsion elements. A method of fitting out such an aircraft is also described.

A gyroscopic orbiter with vertical takeoff and vertical landing capabilities can transition between different functional modes while in-flight. The orbiter typically includes a fuselage, a front boom, a front propulsion unit, a rear boom, and a rear propulsion unit. The front boom is mounted at two pivot points to a bow of the fuselage by the front boom. The rear boom is mounted at two pivot points to a stern of the fuselage by the rear boom. One functional mode is the vertical takeoff and landing mode, wherein the propulsion units are oriented parallel to each other and are directed upward. Another functional mode is the shuttle mode, wherein the propulsion units are oriented at an angle with each other, and the front propulsion unit is directed forward. Another functional mode is the high speed mode, wherein the propulsion units are oriented collinear with a roll axis of the fuselage.