823results about "Power plant type" patented technology

An autonomous flying device achieving a large payload and a long continuous flight time and also accurately adjust position and orientation while flying. The device includes: a main rotor and the like that provide main thrust; a sub rotor and the like that controls the orientation; an engine that generates energy for rotating the main rotor and the like and the sub rotor and the like; and an arithmetic control device that controls rotation of the sub rotor and the like. Also, the main rotor and the like are rotated by being drivingly connected to the engine, whereas the sub rotor and the like are rotated by motors driven by electric power generated from generator and the like operated by the engine. Further, when orientation control to tilt the fuselage is performed, the arithmetic control device increases the output distribution ratio of the sub rotor to above the output distribution ratio of the sub rotor when hovering is performed.

A vertical take-off or landing (VTOL) aircraft is disclosed that includes a propulsion system comprising at least four tilting propulsion assemblies and two or more lifting propulsion assemblies. Actuation of the tilting mechanisms can convert the aircraft from a hover configuration, in which all propellers produce thrust generally upward to counteract the aircraft's weight, to a forward flight configuration, in which the four or more tilting propellers produce thrust generally in a forward direction parallel to the fuselage to overcome drag in cruise flight and the lifting propellers produce no thrust or a level of thrust generally less than that required in the hover configuration.

A triple-flow turbine engine for an aircraft, this turbine engine having a longitudinal axis and including upstream suspension elements which are located in a first plane perpendicular to the axis and are connected or fixed to the gas generator of the turbine engine; downstream suspension elements which are located in a second plane perpendicular to the axis and are connected or fixed to the gas generator; and thrust-absorbing rods which include first ends which are connected or fixed to the gas generator and opposing second ends which are located in a third plane perpendicular to the axis, wherein the first, second and third planes are located at a cold compartment of the gas generator.

A system and a method for controlling a propulsor assembly of an electric aircraft. The system may include an electric motor, wherein the electric motor may include a rotor and a stator, a propulsor driven by the electric motor configured to propel an electric aircraft. The propulsor may include a propeller. The propeller may include a blade. The system may include a cyclic control assembly configured to deflect the blade. The cyclic control assembly may include an actuator and a push rod mechanically connected to the actuator and the propulsor. The system may include a flight controller. The flight controller may be configured to receive sensor datum from at least a sensor communicatively connected to the electric aircraft, generate a deflection command as a function of the sensor datum and actuate the cyclic control assembly as a function of the deflection command.

An autonomous flying device achieving a large payload and a long continuous flight time and also accurately adjust position and orientation while flying. The device includes: a main rotor and the like that provide main thrust; a sub rotor and the like that controls the orientation; an engine that generates energy for rotating the main rotor and the like and the sub rotor and the like; and an arithmetic control device that controls rotation of the sub rotor and the like. Also, the main rotor and the like are rotated by being drivingly connected to the engine, whereas the sub rotor and the like are rotated by motors driven by electric power generated from generator and the like operated by the engine. Further, when orientation control to tilt the fuselage is performed, the arithmetic control device increases the output distribution ratio of the sub rotor to above the output distribution ratio of the sub rotor when hovering is performed.

The present disclosure relates to a reconfigurable battery system and method of operating the same. An example apparatus includes at least one memory, instructions in the apparatus, and processor circuitry to execute the instructions to determine a state of charge of a battery, determine a closed circuit voltage of the battery, determine a value of a parameter based on a ratio of the state of charge and the closed circuit voltage, and control a switch coupled to the battery based on the value of the parameter, the controlling of the switch to either cause the battery to be coupled to a battery string or cause the battery to be disconnected from the battery string.

The invention discloses a battery tank structure of an unmanned aerial vehicle (UAV) and relates to the technologies of battery tank structures of UAVs. The battery tank structure can solve the problem of large space occupation of a battery tank of an existing UAV. A main battery tank body is of a rectangular structure, is placed in the middle of a UAV body, the side wall of the tank body is provided with a plurality of positioning holes, and a battery locking mechanism is fixed by the positioning holes and fixedly connected to a battery after the battery is inserted into the battery tank; byfixing the battery locking mechanism to the positioning holes in different positions, the position of the battery in the whole UAV body is adjusted, and correspondingly the position of the gravity center of the UAV body is adjusted. The structure is mainly used for UAVs.

Examples provide electric propulsion systems; such a system comprising: a. a nacelle comprising an aerodynamic annular housing defining an internal annular volume and defining an inner duct, b. at least one fan stage comprising a fan bearing a plurality of aerodynamic surfaces / blades to generate thrust mounted within a fan shroud, and c. a powertrain comprising at least one electric motor to drive the fan shroud and at least one electrical power source to power the electric motor; and d, wherein i. the fan shroud is disposed within the inner duct, ii. at least part of the powertrain is disposed within the internal annular volume, iii. the at least one electric motor is disposed within the central hub and iv. the hub-to-tip ratio is between 0.1 and 0.4, preferably, between 0.2 and 0.3.

An assembly for a propulsion system of an aircraft includes a gearbox module, an electric motor assembly, a first accessory load assembly, and a propulsor. The gearbox module includes a gear assembly and an output shaft. The gear assembly includes a main gear and a plurality of offset gears. The electric motor assembly includes an electric motor. The electric motor includes a rotor. The rotor is mounted to a first offset gear of the plurality of offset gears. The rotor is configured for rotation about a rotational axis to drive rotation of the gear assembly and the output shaft. The first accessory load assembly includes at least one first accessory load. The at least one first accessory load is mounted to a second offset gear of the plurality of offset gears. The propulsor is coupled to the output shaft.

The invention discloses a performance simulation calculation method for an aviation hybrid electric power system. The method comprises the following steps: firstly, constructing a sub-component calculation model covering an aerodynamic thermal component, a power generation subsystem, a power conversion and transmission subsystem, a battery subsystem, a motor and an aerodynamic system; establishing a reverse logic calculation framework driven by a load side demand, and realizing hierarchical reverse parameter solution from a load end to a power source end; and establishing a power balance equation at a bus end, carrying out iterative solution by adopting a Newton-Rafson algorithm, and realizing dynamic distribution and coordination control of multi-source energy in combination with an energy constraint module. According to the method, through a reverse solution and bus end balance strategy, the simulation calculation efficiency and convergence of the complex aviation hybrid electric power system are remarkably improved, and meanwhile, the adaptability and expandability of the model to different topological architectures and task profiles are enhanced; and an efficient and high-credibility calculation tool is provided for design, performance evaluation and energy management strategy optimization of the hybrid electric propulsion system of the aircraft.

The present invention relates to a vertical take-off and landing (VTOL) aircraft featuring a central fuselage and at least two power units. Each power unit is connected to at least two separate lifting rotors via transmission lines, ensuring that each power unit is linked to at least two rotors diametrically opposed relative to the centre of gravity of the aircraft. The lifting rotors might be driven by power transmission lines of hydrostatic type. The power units on either side might be housed in separate pods which might be incorporated into the main fuselage or arranged as separate pods. These separate pods might also function as floats / pontoons for the aircraft to land, take-off and manoeuvre on water.

A modular battery system includes an array of battery modules arranged in at least one stack. Each battery module includes a plurality of battery cells, a first side having positive and negative receptacles and a second side, that is opposite of the first side, having positive and negative plugs. The receptacles and plugs are configured such that adjacent battery modules in a side-by-side relationship are electrically coupled together via plug and receptacle connections and such that the battery modules are electrically coupled together in parallel. An interconnection electrically couples each stack of battery modules together via plug and receptacle connections with one of the battery modules in each stack such that the stacks of battery modules are electrically coupled together in parallel. Each of the battery modules includes a voltage regulator configured to convert voltage between a battery cell voltage and a bus voltage.

An aerial vehicle connected to or integrated with a work device, including a power source to generate power, a rotor to generate lift used for flight by rotating a propeller, and a power transmission to transmit driving power to the work device.

Embodiments provide an aircraft with one or more tilting fan assemblies that are configured to tilt between a forward flight position and a vertical lift position. The aircraft may also include a plurality of lift fan assemblies for vertical movement. The tilting fan assemblies may be coupled to the fuselage or wings of the aircraft via one or more tilting mechanisms. A control system coupled to the aircraft may control the one or more tilting mechanisms to move the tilting fan assemblies between the forward flight position and the vertical lift position. The tilting fan assemblies may be coupled to one or more support structures that are coupled the fuselage or wings of the aircraft.

An aircraft equipped with a multi-fan propulsion system for controlling flight orientation transitions is provided. In one example aspect, an aircraft includes a fuselage and a pair of wings. The aircraft includes a propulsion system having a first propulsor and a second propulsor each mounted to the fuselage. The first propulsor has a fan positioned primarily above and the second propulsor has a fan positioned primarily below the pair of wings. The aircraft also includes a computing system having one or more processors configured to cause, in response to a demand to change an orientation of the aircraft for a flight orientation transition, the fans of the first and second propulsors to produce different amounts of thrust with respect to one another so that the aircraft performs the flight orientation transition. The thrust differential causes the aircraft to transition between orientations.

An aircraft includes an aircraft heat source; a propulsion system including an electric propulsion engine, the electric propulsion engine including an electric motor and a fan rotatable by the electric motor, the electric propulsion engine further defining a fan air flowpath; a thermal management system including a heat source exchanger in thermal communication with the aircraft heat source, a heat sink exchanger in thermal communication with the fan air flowpath of the electric propulsion engine, and a thermal distribution bus extending from the heat source exchanger to the heat sink exchanger; and a control system operably connected to the thermal management system for selectively thermally coupling the heat sink exchanger with the heat source exchanger.

An apparatus for aircraft propulsion includes a propeller and a hybrid-cooled electric engine mounted to a support structure and configured to rotate the propeller. The electric engine is located in an enclosure, with a first heat transfer element thermally coupled to the electric engine via a fluid flow path, partially located outside the enclosure. A second heat transfer element, integral to or thermally coupled to the electric engine, provides air cooling. Both heat transfer elements are housed within the aircraft's support structure. At least one air inlet on the upper side of the support structure receives propeller downwash. A first cooling path directs a portion of the downwash from the air inlet to the first heat transfer element, and a first air outlet exhausts the downwash.

A VTOL aircraft (300, 600) includes a lift propeller (312, 612) configured to provide lift during take-off, landing, and hover operations. The lift propellers (312, 612) are configured to stow close to the boom surface (323, 623) when not in use, for example, during a cruise flight, to minimize drag around the surface of the propellers (312, 612). The lift propellers (312, 612) are vertically displaceable relative to the booms (322, 622) when switched to a lift configuration to provide a greater spacing from the booms (322, 622) to improve lift efficiency and reduce unwanted vibrations. The lift propeller (312, 612) may have a fail-safe configuration that allows the propeller (312, 612) to passively move between a lift configuration and a stowed configuration, and to fully rotate even when fully retracted.

Systems, apparatus and methods are provided for securing firmware and mission data in memory of an unmanned aerial vehicle. The unmanned aerial vehicle may include a central body and at least one electric motor. The central body may include a flight controller with a processor, a firmware memory configured to receive and store firmware, and circuit board with a fixed portion and a removable portion with a programming interface connector providing access through a signal connection to a firmware memory operatively connected to the flight controller. The central body may further include a volatile mission memory configured to receive mission data and automatically be erased upon power interruption.

The application is suitable for the field of flying cars, and provides a vertical take-off and landing flying car, which comprises a car body, a main lift assembly arranged on the car body and placed inside the car body, a folding wing assembly arranged on the car body and placed on the top of the car body, and a back thrust power assembly arranged on the car body and placed on the top of the tail end of the car body. The application aims to solve the technical problems in the prior art that one kind of flying car is vertically taken off and landed, but does not have a fixed-wing flight mode, provides lift by a rotor, is more like a variant of a helicopter, has a small flight speed, a short flight range, and a complex control; another kind of flying car is taken off by sliding, has a main wing for generating lift, but has a long sliding distance for taking off and landing, has a high requirement for a landing site, and greatly limits the use scene.

An apparatus for aircraft propulsion includes a propeller and an electric engine mounted to a support structure and configured to rotate the propeller. The electric engine is located in an enclosure, with a heat transfer element thermally coupled to the electric engine, An air inlet is located at an upper side of the support structure and the air inlet is configured to receive downwash from the propeller. An air outlet is located at the upper side of the support structure, and the air outlet is configured to exhaust the downwash from the heat transfer element.

A system for an electrically powered aircraft can comprise: a truss configured to couple to a structure of an aircraft; a battery system supported by the truss, the battery system including a first array of battery modules and a second array of battery modules, the first array of battery modules operably coupled to the second array of battery modules: and an exhaust system including a first exhaust manifold, a second exhaust manifold, and a common exhaust manifold, the first exhaust manifold coupled to each battery module in the first array of battery modules, the second exhaust manifold coupled to each battery module in the second array of battery modules, and the first exhaust manifold and the second exhaust manifold each coupled to the common exhaust manifold.

One embodiment is an aircraft including a fuselage; a wing connected to the fuselage; first and second booms connected to the wing on opposite sides of the fuselage; first and second forward propulsion systems attached to forward ends of the first and second booms; first and second aft propulsion systems fixedly attached proximate aft ends of the first and second booms; first and second wing-mounted propulsion systems connected to outboard ends of wings; and first and second wing tips fixedly connected to outboard sides of the first and second wing-mounted propulsion systems; wherein the first and second wing-mounted propulsion systems and the first and second wing tips are collectively tiltable between a first position when the aircraft is in a hover mode and a second position when the aircraft is in a cruise mode.

An air-moving device may include an aerodynamic stator. The aerodynamic stator may be positioned forward of a motor of the air-moving device and aftward of an aerodynamic rotor of the air-moving device. A control unit may be integrated in and in thermal communication with the aerodynamic stator. The aerodynamic stator may transfer heat from the control unit to thermally conductive stator vanes of the aerodynamic stator. An airflow generated by the aerodynamic rotor may facilitate heat dissipation from the thermally conductive stator vanes. The aerodynamic stator may include electrically conductive stator vanes. The electrically conductive stator vanes may provide at least one of power or control signaling to the control unit.

An aircraft hybrid electrical propulsion (HEP) system includes an electrical system configured to deliver power to a plurality of electrical loads, a propulsion system configured to generate thrust in response to an input power, and an HEP controller in signal communication with the electrical system and the propulsion system. The HEP controller is configured to monitor a load demand of at least one electrical load among the plurality of electrical loads and to actively control the input power to actively control the thrust in response to changes in the load demand.

To provide a power supply system, an aircraft, and a power supply system control method that makes the remaining capacity of each battery approximately equal.SOLUTION: A power supply system 26 includes a power converter 36 that converts the power of any one of a plurality of batteries included in a high voltage circuit 32 and supplies it to one or more second loads 38, a switch 34 capable of selectively connecting any one of the plurality of batteries to the power converter, and a controller 70 that compares the remaining capacities SOC of the batteries, and controls the switch so as to connect a battery with the highest remaining capacity and the power converter when the difference between the highest remaining capacity and the lowest remaining capacity is greater than a predetermined threshold.SELECTED DRAWING: Figure 3

Methods, apparatus, systems, and articles of manufacture are disclosed for a hydrogen fuel distribution system (100) with integrated thermal management. An example fuel distribution system (100) with integrated thermal management comprises a hydrogen fuel distributor (402) including a first tank (406) and a second tank (408). The hydrogen fuel distributor (402) circulates hydrogen fuel to a first fuel cell (606) and a second fuel cell (608). A coolant distributor (702) includes a heat exchanger (416). The coolant distributor (702) regulates a temperature of the fuel distribution system (100). A compressed air distributor (706) provides compressed air from a propulsor (730) to the first fuel cell (606) and the second fuel cell (608). An electrical distributor (602) transports electricity from the first fuel cell (606) and the second fuel cell (608) to an electric motor (726). The electric motor (726) drives the propulsor (730).

Aircraft turbomachine (10), comprising: - an engine (12), - an annular panel (16) extending around the engine (12), - a support structure (14) for the engine (12), - an annular gas flow channel (18) between the panel (16) and the structure (14), - a strut (20) connecting the panel (16) to the structure (14), and - a device (22) for attaching the strut (20) to the panel (16), this device (22) comprising a fitting (24) which is attached by first screws (26) to the strut (20) and by second screws (28) to the panel (16), the first screws (26) being screwed into nuts (32) which are mounted inside at least one internal cavity (40) of the strut (20), this internal cavity (40) being located radially at the level of said vein (18). Figure for the abbreviation: Figure 3

The invention relates to an electric propulsion system comprising: - a propeller (32) having an axis of rotation (A32), - at least two electric motors (34) offset from each other along a longitudinal direction parallel to the axis of rotation (A32) of the propeller (32), - a main drive shaft (42) parallel to the axis of rotation (A32) of the propeller (32) to which the electric motors (34) are coupled, - a transmission system (36) comprising an output shaft (38) connected to the propeller (32) and an input shaft (38') connected to the main drive shaft (42). This solution simplifies the design of the transmission system and reduces the cross-section of the electric propulsion system at the location of the electric motors. The invention also relates to an aircraft comprising at least one such electric propulsion system.