Abstract: A control system for a hybrid electric powerplant of an aircraft can include a throttle controller configured to receive one or more power settings and to output a heat engine setting and an electric motor setting, a heat engine controller operatively connected to the throttle controller. The heat engine controller can be configured to receive the heat engine setting and to control a heat engine system as a function of the heat engine setting to control torque output by a heat engine. The system can include a heat engine protection module that is part of or connected to the heat engine controller and configured to provide one or more protection commands to directly control one or more heat engine protection systems. The system can include an electric motor controller operatively connected to the throttle controller.

Abstract: Landing apparatuses for unmanned aerial vehicles are provided herein. An example UAV includes a frame; a propeller rotatably coupled to the frame; and a landing guard armature extending from the frame. A terminal end of the landing guard armature extends beyond a propeller radius of the propeller. The landing guard armature has a surface area that is sized to promote airflow around the landing guard armature.

Abstract: Provided is a multicopter providing a high level of freedom in compact design, and consuming a relatively small amount of energy. The multicopter (10) includes a machine body (12), an N number of first lift generators (30) arranged on a first concentric circle (C1) centered substantially around a gravitational center (G) of the machine body (12) and in a front part and a rear part of the machine body in a bilateral symmetry, and an M number of second lift generators (70) arranged on a second concentric circle (C2) centered substantially around the gravitational center (G) of the machine body (12) and having a larger diameter than the first concentric circle (C1), and in a front part and a rear part of the machine body (12) on a central axial line (X) extending in a fore and aft direction of the machine body (12), N being greater than M.

Abstract: A system and method for an aircraft evacuation system with hydrogen inflation is disclosed. The system includes an aircraft having an integrated hydrogen-electric engine. A fuel cell stack for powering an aircraft motor of the integrated hydrogen-electric engine. A hydrogen fuel source in fluid communication with the fuel cell stack, the hydrogen fuel source containing hydrogen. An inflatable slide and a pump operably coupled with the hydrogen fuel source and the inflatable slide to selectively pump the hydrogen to the inflatable slide for inflating the inflatable slide.

Abstract: An aircraft includes an airframe, and a coaxial main rotor assembly including a static mast and an upper rotor assembly and a lower rotor assembly rotatable about a main rotor axis defined by the static mast. The upper rotor assembly and the lower rotor assembly are independently rotatable about the static mast. A propulsion system includes at least one propulsion source for directly driving at least one of the upper rotor assembly and the lower rotor assembly and a flight control system is operably coupled to the propulsion system. The flight control system is operable to independently control a rotational speed of the upper rotor assembly and the lower rotor assembly relative to the static mast.

Abstract: Embodiments relate to reconfigurable unmanned vehicles (100). Such vehicles (100) comprise a fuselage (102) presenting a bay (118) for receiving a plurality of components (120-128), each of the plurality of components (120-128) relating to a respective entity for at least one of flight control or operation of the unmanned vehicle (100). The bay (118) comprising a bus to support communications between at least two of the plurality of components (120-128), the plurality of components (120-128) comprising a controller to determine a configuration or presence of one or more than one component of the plurality of components (120-128) when coupled to the bus.

Abstract: A hydrogen storage system comprises a hydrogen tank and a system for controlling hydrogen evaporation in the hydrogen tank. This control system comprises a hydrogen discharge pipe connected to the hydrogen tank, on the one hand, and to a controllable valve, on the other hand, as well as a processing unit configured to control the valve as a function of the pressure in the tank. The hydrogen storage system further comprises a fuel cell permanently connected to the hydrogen tank and the processing unit being electrically powered by the fuel cell.

Abstract: An aircraft including: a wing having a leading edge; an electric motor mounted on the wing; and an electric power cable in the wing for providing power to the electric motor, wherein the electric power cable runs through a duct in the wing; the duct comprises an air inlet and an air outlet; and the electric power cable is cooled by air passing through the duct between the air inlet and the air outlet.

Abstract: An aircraft hybrid propulsion system comprises an internal combustion engine, an electric motor a propulsor and a combining gearbox. The internal combustion engine is coupled to a first input of the combining gearbox, the electric motor is coupled to a second input of the combining gearbox, and the propulsor is coupled to an output of the combining gearbox, such that the propulsor is driveable in use by either or both of the internal combustion engine and the electric motor. Each of the internal combustion engine and the electric motor is coupled to its respective input by a respective clutch.

Abstract: An aircraft propulsion system capable of improving the restartability of an engine is provided. An aircraft propulsion system includes: a plurality of engines (60-1 and 60-2); an electric generator connected to engine shafts (63-1 and 63-2); a plurality of electric motors; a plurality of rotors; and a controller (100) which controls an operation state of the plurality of engines (60-1 and 60-2). The controller (100) causes a part of the plurality of engines (60-1 and 60-2) to operate while stopping the remaining engine when a flight state is in a first state. The plurality of engines (60-1 and 60-2) include circulators (65-1 and 65-2). A plurality of communication pipes (68 and 69) which communicate the circulator (65-1) in the first engine (60-1) among the plurality of engines (60-1 and 60-2) with the circulator (65-2) in the second engine (60-2) are provided.

Abstract: An auxiliary power unit (APU) system for an aircraft can include a fuel consuming system configured to generate and output electrical energy for use by one or more aircraft systems and a battery system configured to output stored electrical energy for use by the one or more aircraft systems. The fuel consuming system can include a fuel consuming APU, and an APU generator operatively connected to the fuel consuming APU and configured to convert APU motion into APU generator alternating current (AC). The system can include a high voltage AC line operatively connected to the generator. The battery system can include a battery configured to output battery direct current (DC).

Abstract: A fuel containment system for an aircraft is provided. The fuel containment system comprises an upper fuel barrier under a cabin floor, an aft wheel well bulkhead, an aft fuel barrier opposite the aft wheel well bulkhead, a lower fuel barrier associated with a cargo floor, and a lower fuselage skin panel. A fuel tank is created by the upper fuel barrier, the aft wheel well bulkhead, the aft fuel barrier, the lower fuel barrier, and the lower fuselage skin. The fuel tank is integrated into the aircraft and existing structural components are sealed to prevent fuel from leaking out of the integrated fuel tank.

Abstract: A method of reducing the noise generated in-flight by a vortex wake caused by each first blade of a main rotor of a hybrid helicopter. The hybrid helicopter includes a main rotor, at least two wings and at least one propeller. The method enables a stabilized flight phase on the level or with a non-zero aerodynamic slope to be implemented by determining a first value of the pitch of the second blades of each propeller and an angle of incidence of the main rotor as function of the flight conditions, then by applying the first pitch value to each propeller and by applying the angle of incidence to the main rotor so as to direct the vortex wake to limit the noisy interactions between the vortex wake and the other first blades and/or the second blades.

Abstract: An aircraft comprising a wing having a spanwise lift distribution extending from a root to a tip, the lift distribution defining an inboard region defining a positive lift contribution, an outboard region defining a negative lift contribution, and an intermediate region defining a neutral lift contribution, the neutral region being spaced from the tip and from the root. A propulsion system is provided, comprising a wing mounted propulsor. The wing mounted propulsor has a rotational axis (x) positioned substantially at a span of the wing where a value of ?Lift/?Span is at a maximum for the span of the wing, and may be located at the intermediate region along the span of the wing.

Abstract: 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.

Abstract: A rotorcraft has a horizontal stabilizer movable about an axis of rotation and a horizontal stabilizer control system configured to control the horizontal stabilizer to at least one of move the rotorcraft into a minimum drag position, maintain the aircraft in a minimum drag position, efficiently achieve a maneuver, enter efficient autorotation, and maintain efficient autorotation.

Abstract: A leading edge structure for a flow control system of an aircraft is disclosed having a leading edge panel that surrounds a plenum, wherein the leading edge panel has a first side portion, a second side portion opposite the first side portion, an inner surface facing the plenum and an outer surface in contact with an ambient flow, and wherein the leading edge panel comprises a plurality of micro pores forming a fluid connection between the plenum and the ambient flow, wherein the plenum is connected to an air outlet arrangement configured for causing an underpressure in the plenum, so that air from the ambient flow is drawn through the micro pores into the plenum and from there discharged through the air outlet arrangement into the ambient flow.

Abstract: A propulsion system for providing propulsion of an aircraft includes a plurality of electric motors coupled with a rotor of the aircraft to drive the rotor and a propulsion motor control. The propulsion motor control includes at least one processor electrically connected with at least one electric motor to actuate the at least one electric motor and at least one battery electrically connected with the at least one processor and at least one electric motor to provide power to the at least one processor and the at least one electric motor. The propulsion motor control actuates the plurality of electric motors based on a desired torque level to drive the rotor to provide propulsion of the aircraft.

Abstract: An electric aircraft includes rotors for providing lift for vertical take-off and landing of the aircraft, proprotors that are tiltable between lift configurations for providing lift for vertical take-off and landing of the aircraft and propulsion configurations for providing forward thrust to the aircraft, a first battery pack for powering a first rotor and a first proprotor, a second battery pack for powering a second rotor and a second proprotor, a first electric power bus electrically connecting the first battery pack to the first rotor and proprotor, and a second electric power bus electrically connecting the second battery pack to the second rotor and proprotor, wherein the second electric power bus is electrically isolated from the first electric power bus.

Abstract: A method and system for attaching a vertical stabilizer to an aircraft fuselage using a clevis system is disclosed. A composite skin is installed over the aircraft fuselage. The composite skin has apertures for receiving a plurality of clevises in a clevis system. The plurality of clevises is inserted through the apertures in the composite skin. Each of the plurality of clevises is secured to a frame member in the aircraft fuselage. The vertical stabilizer has a multi-spar box connected with a base rib assembly having a plurality of lugs. The base rib assembly of the vertical stabilizer is engaged with the clevis system. The plurality of lugs in the base rib assembly is secured to the plurality of clevises in the clevis system. Attachment of the vertical stabilizer to the aircraft fuselage, as well as subsequent inspection, may be performed from outside the aircraft.