A hybrid power or thrust generator including at least one thermodynamic turbine engine and at least one electric power generator. An electric power generator is electrically connected to at least one electric motor, mechanically and rotatably coupled to one or more rotating portions of the thermodynamic turbine engine, and the electric power generator operates simultaneously with the thermodynamic engine so as to supply the electric motor(s) to reduce the power drawn from the turbines of the thermodynamic turbine engine in operation. The electric power generator may comprise a thermoacoustic engine driving a linear electric alternator. The generator is advantageously implemented in a vehicle, such as an aircraft.

A first DC power supply section of a DC stabilization power supply system includes at least a generator and a PWM converter. A second DC power supply section of the DC stabilization power supply system includes at least a battery and a boost converter. Each of these DC power supply sections performs a power supply stabilization operation for supplying DC power to an electric system and absorbing regenerative power generated from an actuator or the like. A PWM converter controls adjustment of the amount of the power supply stabilization operation to be performed by the DC power supply section and the amount of the power stabilization operation to be performed by the DC power supply section, by changing a target voltage value of a DC bus.

An aircraft conditioning system including a fuel tank, including an outlet connected to an Auxiliary Power Unit (APU) fuel inlet, configured to maintain a pressurized gas fuel, a pressure regulator, downstream of the fuel tank, configured to regulate a pressure of the pressurized gas fuel injected to the APU inlet, a first temperature sensor downstream of the pressure regulator and a second temperature sensor upstream of an engine, a first pressure sensor inside the fuel tank, a second pressure sensor downstream of the pressure regulator and a third pressure sensor at an engine inlet, and a controller connected to the pressure sensors and the temperature sensors, the controller receiving temperature and pressure values from the pressure sensors and from the temperature sensors and actuating on at least one of the pressure regulator or the injection position of the pressurized gas fuel inside a plurality of engine combustion chambers.

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.

An exemplary boundary layer ingestion engine includes a gas generator, a turbine fluidly connected to the gas generator, and a fan mechanically linked to the turbine via at least one shaft. The linkage is configured such that rotation of the turbine is translated to the fan. The boundary layer ingestion engine further includes an exhaust duct fluidly connected to an outlet of the turbine. The exhaust duct is positioned radially inward of the fan.

The invention relates to an aircraft incorporating an enhanced power unit for generating electric, pneumatic and/or hydraulic power for the aircraft during all stages of the aircraft operation. The power unit (1) comprises: a heat engine (14) with a drive shaft (2) and a combustion gases exhaust (7). The power unit (1) also includes a Rankine cycle system (12) for recovering thermal energy from a heat source of the power unit (1) for the assistance of the heat engine (14). The heat source for the Rankine cycle system can be taken from the exhaust gases of the heat engine, from the oil coolant circuit of the heat engine or from the output of a compressor driven by the heat engine. Preferably, the aircraft cabin air is reused as a source of oxygen for the combustion. The invention reduces bleed air extraction from the aircraft main engines thereby reducing fuel consumption.

A vertical take-off and loading (VTOL) rotary aircraft or helicopter has eight propellers in a quad propeller arm configuration where each propeller arm has two counter-rotating propellers. Folding propeller arms are designed to allow storage in a single car sized garage. Each propeller may be powered by a three-phase alternating current motor. The main power plant for the aircraft is a gas combustion engine that generates electricity. If the gas engine fails, a battery backup system will safely bring the aircraft down for a controlled landing. The direct current bus is redundant in that even with a gas combustion engine failure the direct current bus battery pack will safely bring down the aircraft. Various embodiments of this invention may also include a landing gear crumple zone designed to soften a hard landing.

A mounting assembly for mounting a portion of an engine exhaust duct in an opening of an airframe panel aircraft, including a peripheral flange disposed on the portion of the exhaust duct, and at least one mounting member having an outer peripheral flange portion for connection to the airframe panel and having an inner peripheral flange portion with at least one compliant isolator member to receive the peripheral flange of the exhaust duct in a manner that suspends the exhaust duct in the opening away from the airframe panel and accommodates thermal expansion and contraction of the exhaust duct relative to the airframe panel. Embodiments include a method of mounting an engine exhaust duct in an opening of an outer airframe panel of an aircraft.

A suspension system for an aircraft auxiliary power unit located in a fuselage structure, the system including: struts (10), auxiliary power unit attachment brackets (51,52) for connecting the strut (10) to the auxiliary power unit (30), vibration isolators (5) for joining the struts (10) and the auxiliary power unit attachment bracket (51, 52), a cone-bolt (1) attached to the auxiliary power unit attachment brackets (51,52) and having a longitudinal threaded hollow, an inner bolt (2) partially located within the hollow of the cone-bolt (1) and threaded to it (1), an outer bolt (3) having a longitudinal through-hole and partially located within the hollow of the cone-bolt (1) and including an external thread that engages the thread of the cone-bolt (1), the inner bolt (2) extending across the hollow of said outer bolt (3).

The present invention relates to a drive device assembly having a plurality of drive devices and an aircraft for which drive devices and loads can be increased unlimitedly. The drive devices are arranged in a triangle so as to form a triangular drive device module. In a preferred assembling method, the spatial distances between adjacent drive devices are equal; a drive device module arranged in a triangle is taken as a basis, and is mapped and arranged and assembled towards the spatial direction so as to form a drive device assembly constituted by the multiple drive devices. The present application improves the aircraft, especially in terms of the traditional structure layout of the aircraft driven by multiple drive devices, such that the drive devices in the drive system can be increased unlimitedly, thus improving the flight performance and loading capacity of the aircraft.