1202results about "Power plant cooling arrangments" patented technology

The aircraft includes a fuselage, a front wing and a rear wing provided to extend laterally from the fuselage for generating lift during cruise, a boom extending in a front-back direction supported by these wings to be spaced apart from the fuselage, a VTOL rotor supported on the boom and having one or more blades for generating thrust in a vertical direction during take-off and landing, a cooling system for cooling the VTOL rotor by using a radiator arranged within the boom, and an airflow guide structure provided in a part of the boom for guiding airflow generated by rotation of the blades to the radiator, and having an inlet provided on a side that faces toward a rotational direction of the blades relative to a rotational axis of the VTOL rotor, of a surface of the boom positioned below a plane of rotation of the blades.

An aircraft propulsion assembly that has a nacelle with a front opening and a rear opening, a propeller driven by an electric motor, a fuel cell supplying the electric motor with electric current, an inner channel arranged in the nacelle between the front opening and the rear opening. The electric motor and the fuel cell are disposed in the inner channel. A motorized fan is arranged at the rear part so as to draw in the air present in the inner channel and expel it to the outside through the rear opening. A control unit controls the motorized fan. With such an arrangement, the outside air cools the inside of the nacelle and the dihydrogen is evacuated.

An aircraft system is provided. The aircraft system includes: a gas turbine engine; a fuel system having a fuel tank and a fuel delivery system fluidly connecting the fuel tank to the gas turbine engine; an electric drive assembly including a component module; and a thermal management system including: a thermal fluid loop; a heat source heat exchanger in thermal communication with the component module and the thermal fluid loop; an airflow heat sink heat exchanger in thermal communication with the thermal fluid loop and further configured to be in thermal communication with a cooling airflow during operation of the gas turbine engine; and a fuel heat sink heat exchanger in selective thermal communication with the thermal fluid loop, the fuel tank, or both, the fuel heat sink heat exchanger in fluid communication with the fuel tank independently of the fuel delivery system during at least a first operating condition.

A thermal management system defines a thermal management system flowpath to provide a flow of a fluid to an electric machine and a power electronics assembly electrically connected to the electric machine. The thermal management system includes a first heat exchanger thermally connected to the thermal management system flowpath and to the electric machine, and a second heat exchanger thermally connected to the thermal management system flowpath downstream of the first heat exchanger. The second heat exchanger is thermally connected to the power electronics assembly.

The invention provides an electric engine, a control method, a related device, a motor controller and an aircraft, and relates to the technical field of aircrafts. The electric engine comprises a power motor, a heat dissipation system and a control module, a cavity is formed in the power motor, the control module is integrated in the cavity, the space occupied by the electric engine is reduced, and therefore the size of the aircraft is reduced; the control module comprises a plurality of integrated driving modules, part of the driving modules are electrically connected with the power motor so as to control the power motor, and part of the driving modules are electrically connected with a heat dissipation motor in the heat dissipation system so as to control the heat dissipation motor. The performance and efficiency of the electric engine are improved through unified integrated control over the power motor and the heat dissipation motor. And through a unified control architecture, the system design is simplified, the development and maintenance cost is reduced, and the maintainability of the electric engine is improved.

An assembly for an aircraft includes a wing assembly, a plurality of underwing heat exchangers, and a mounting assembly. The wing assembly includes a wing body. The plurality of underwing heat exchangers are arranged along the wing body in a longitudinal direction. The plurality of underwing heat exchangers includes a first underwing heat exchanger and a second underwing heat exchanger. The first underwing heat exchanger is longitudinally adjacent the second underwing heat exchanger. The mounting assembly includes a first heat exchanger bracket, a second heat exchanger bracket, and a first support bracket. The first heat exchanger bracket is fixedly mounted to the first underwing heat exchanger. The second heat exchanger bracket is fixedly mounted to the second underwing heat exchanger. The first heat exchanger bracket and the second heat exchanger bracket are pivotably mounted to the first support bracket. The first support bracket is fixedly mounted to the wing body.

The invention relates to an aircraft electric drive oil cooling structure which comprises an outer rotor assembly, a stator assembly, a controller assembly, an oil pump and a heat dissipation assembly, the stator assembly comprises a stator shell and a winding, a sealing cover is further fixed to the outer side of the winding, and a cooling cavity is formed between the sealing cover and the stator shell; a bottom plate is further arranged at the bottom of the stator shell, the controller assembly is fixed in a mounting cavity formed by the bottom plate and the stator shell and comprises a circuit board and a cooling plate which are close to each other, the circuit board is connected with the winding, an oil duct A is further arranged on the bottom plate, an oil inlet pipe and an oil outlet pipe are arranged on the cooling plate, and an oil duct B communicated with a cooling cavity is further arranged on the bottom plate; the heat dissipation assembly comprises a heat dissipation fin set and a fan assembly, the heat dissipation fin set and the oil pump are both fixed to the lower portion of the bottom plate, heat dissipation fins form a ring and are arranged outside the oil pump in a surrounding mode, the heat dissipation fin set, the oil pump and the bottom plate are communicated with one another, and the fan assembly is fixed to the lower portion of the heat dissipation fin set. Efficient, light-weight, high-integration-level and reliable heat dissipation of the electric drive of the aircraft is achieved.

The invention relates to an electric duct driving system integrated with a heat dissipation function and an aircraft, the aircraft comprises the electric duct driving system, the electric duct driving system comprises a casing, an electric driving assembly and a heat dissipation pipeline, a fan is arranged in the casing, and the electric driving assembly is in transmission connection with the fan and used for driving fan blades of the fan to rotate; the heat dissipation pipeline is arranged on the outer surface of the casing and provided with a cooling liquid inlet and a cooling liquid outlet, one part of the heat dissipation pipeline is connected to the outer surface of the casing, one end of the heat dissipation pipeline is connected to the cooling liquid outlet, and the other end of the heat dissipation pipeline is connected to the cooling liquid inlet, so that the electric duct driving system can utilize structural parts (namely the casing and the fan) of the electric duct driving system as components for heat dissipation; therefore, a heat management system does not need to be installed, the structure of the electric duct driving system is greatly simplified, the risks of system weight increment and part failure caused by the adoption of a centralized liquid cooling system are effectively reduced, the energy consumption of the whole machine is also effectively reduced, and the cruising ability of the aircraft is improved.

The invention belongs to the technical field of aviation electric propulsion, and relates to a hydrogen fuel cell superconducting hybrid electric propulsion system based on hydrogen and helium heat exchange cooling, which is suitable for megawatt level branch aircrafts and comprises a liquid hydrogen storage tank, a hydrogen and helium heat exchanger, a hydrogen fuel cell, a lithium battery, an inverter, a superconducting motor, a propeller and other core components. The liquid hydrogen storage tank vaporizes liquid hydrogen through the micro-channel plate type hydrogen-helium heat exchanger and stably supplies the vaporized liquid hydrogen to the hydrogen fuel cell to generate electric energy; helium is used for a low-temperature cooling loop and cooling key assemblies such as a superconducting motor, a low-temperature inverter, a superconducting cable and a hydrogen fuel cell, and it is ensured that the key assemblies stably operate at the optimal working temperature. The hydrogen fuel cell and the lithium battery cooperatively provide electric energy for the superconducting electric propulsion system, the integrated energy management controller dynamically adjusts the output power of the fuel cell and the lithium battery by adopting a state machine strategy, and accurate monitoring and optimal distribution of working parameters of all components are achieved. The optimal heat balance and power output are ensured to be realized in different flight stages such as takeoff, cruise and landing; according to the hydrogen fuel cell hybrid electric propulsion system based on hydrogen and helium heat exchange cooling, the power generation efficiency and the operation stability of the hydrogen fuel cell are remarkably improved, the overall energy conversion efficiency is improved, and the high-power and high-efficiency power extraction requirements of megawatt aircrafts are fully met.

The invention belongs to the technical field of aircraft design, and particularly relates to an integrated tilting power system integrated with hydrogen fuel cell motor electric control and an aircraft. Comprising a tilting mechanism connected with a machine body; the integrated motor and controller is fixedly connected with the tilting mechanism and located at the end, close to the upstream of the incoming flow, of the tilting mechanism; the tilting rotor wing is mounted at one end, close to the incoming flow upstream, of the integrated motor and controller; the hydrogen fuel cell is fixedly connected with the tilting mechanism and is positioned at one end, close to the downstream of the incoming flow, of the tilting mechanism; the nacelle shell wraps the tilting mechanism, the integrated motor, the controller and the hydrogen fuel cell; the nacelle outer heat dissipation fins wrap the nacelle in the circumferential direction; and the tilting mechanism, the integrated motor and controller and the liquid cooling radiating pipeline of the hydrogen fuel cell exchange heat with the radiating fins outside the nacelle.

The system can include an on-board thermal management subsystem. The system 100 can optionally include an off-board (extravehicular) infrastructure subsystem. The on-board thermal management subsystem can include: a battery pack, one or more fluid loops, and an air manifold. The system 100 can additionally or alternatively include any other suitable components.

An aircraft system includes an electrified powertrain having a first power module and a second power module. The aircraft system further includes a thermal management system having a thermal fluid loop for conveying a thermal fluid. The first power module and the second power module are in thermal communication with the thermal fluid loop. The thermal management system further includes a heat exchanger in thermal communication with the thermal fluid loop upstream of the first power module and the second power module such that the thermal fluid from the heat exchanger is partitioned between the first power module and the second power module.

A fluid distribution system for an engine including a fluid chamber. The fluid distribution system may include a first control opening in fluid communication with the fluid chamber. A rotor coupled to the fluid chamber may be configured to rotate the chamber about the axis to cause the fluid chamber to distribute fluid through the first control opening. A first channel coupled to the first control opening may be configured to receive the fluid through the first control opening. A second control opening may be formed in the first channel. A relief opening in the first channel may be configured to expose the first channel to an ambient pressure outside the first channel. The relief opening may be located downstream of the first control opening in a fluid flow direction of the first channel and upstream of the second control opening in the fluid flow direction of the first channel.

Systems and methods for controlling an electric vertical take-off and landing (eVTOL) aircraft are disclosed. In one embodiment, a system includes a processor, a first manipulator communicatively coupled with the processor, the first manipulator configured to accept longitudinal and lateral linear movements as manual inputs and provide corresponding signals to the processor, and a second manipulator communicatively coupled with the processor, the second manipulator is configured to accept longitudinal and lateral linear movements as manual inputs and provide corresponding signals to the processor, wherein the processor is configured to control the heading of the aircraft using the signals received from the second manipulator corresponding to the lateral linear movement of the second manipulator. Some embodiments may additionally include at least one sensor and a thumb rocker for each manipulator.

The utility model discloses an eVTOL aircraft air cooling motor installation structure which comprises a motor base, a motor is installed on the motor base, a wind scooper is sleeved on the outer side of the motor, a skin covers the upper end of the motor base, the upper end of the wind scooper is connected with the skin, heat dissipation fins are arranged on the surface of the motor, and the heat dissipation fins are connected with the wind scooper. A gap of an annular structure is arranged between the wind scooper and the radiating fins. According to the structure, the heat dissipation effect is good, the connecting turned-over edge of the wind scooper is directly turned over to the skin and connected with the skin, no gap exists between the connecting turned-over edge and the skin, heat dissipation airflow can effectively flow into the space between the wind scooper and the motor, the flow speed is increased, heat dissipation is conducted on the motor, and the heat dissipation effect is good.

The invention relates to the technical field of hovercar oil coolers, in particular to a hovercar oil cooler multi-physics field cooperative control method based on reinforcement learning. A three-dimensional continuous action space (fin opening, magnetofluid flow velocity and phase change material activation proportion) is designed in combination with a reinforcement learning algorithm, and collaborative optimization of magnetofluid heat conduction, phase change material heat storage and air heat dissipation is realized in control of the oil cooler of the hovercar for the first time. According to the design, the limitation of traditional single physical field control is broken through, the thermal management efficiency of the system in cross-airspace and multi-working-condition scenes is remarkably improved, and meanwhile energy consumption is saved and the service life of equipment is prolonged.

A hybrid-electric aircraft propulsion system includes a gas turbine engine and an electrical assembly. The gas turbine engine includes a first rotational assembly. The first rotational assembly is rotatable about a rotational axis of the gas turbine engine. The first rotational assembly includes a first shaft, a bladed first compressor rotor, and a bladed first turbine rotor. The first shaft interconnects the bladed first compressor rotor and the bladed first turbine rotor. The electrical assembly includes a first motor-generator, a first motor control unit, a motor-generator (MG) cooling system, and a motor control unit (MCU) cooling system. The first motor-generator is coupled to the first shaft. The first motor control unit is electrically connected to the first motor-generator. The MG cooling system is connected in fluid communication with the first motor-generator. The MCU cooling system is connected in fluid communication with the first motor control unit. The MCU cooling system is independent of the MG cooling system.

In an embodiment a vehicle includes a boom having a first opening at a first surface and a second opening at an opposite second surface, a channel extending in a first direction between the first and second openings and connecting the first opening with the second opening and a heat exchanger arranged in the channel, wherein the heat exchanger is tilted with respect to the first direction.

An electric propulsion system includes a fan rotor assembly having a first fan portion and a second fan portion. A fan cowling defines a first flow passage and a second flow passage. An electric drive mechanism drives the fan rotor assembly. An electric drive cooling system includes a liquid coolant that provides cooling to at least a part of the electric drive mechanism, and a heat exchanger through which the liquid coolant flows is arranged within the second flow passage. The first fan portion provides a first flow of air through the first flow passage to provide at least one of a lifting force or a thrust force to the electric propulsion system, and the second fan portion provides a second flow of air through the second flow passage to provide a flow of cooling air therethrough that is in thermal communication with the heat exchanger.

The invention relates to an electric propulsion heat dissipation system, a power device and an aircraft, and relates to the technical field of aircrafts. The electric propulsion cooling system comprises propellers, arms, power motors and radiators, the electric propulsion cooling system is fixedly connected to an aircraft body structure of the aircraft through the arms, the propellers are arranged at the output ends of the power motors, and the power motors are installed on the arms and can drive the propellers to rotate around the axes of the power motors. The radiator dissipates heat of the power motor through cooling liquid circulation, the vehicle arms are provided with heat dissipation grooves and air outlets, the radiator is installed in the heat dissipation grooves, and orthographic projections of blades of the propeller on the corresponding surfaces of the vehicle arms at least partially cover the radiator, so that airflow generated by the propeller can penetrate through the radiator and then enter the external space through the air outlets. According to the electric propulsion heat dissipation system, the power device and the aircraft, the problem that the lightweight design of the aircraft is affected due to the fact that a current cooling system is large in size and weight is solved.

A hybrid-electric propulsion system includes a gas turbine engine comprising a high pressure system and a low pressure system, an electric machine coupled to one of the high pressure system or the low pressure system, and a thermal management system defining one or more thermal management system flowpaths to provide one or more heat exchange fluids to the electric machine. A controller collects one or more signals from one or more sensing nodes connected to at least one of the thermal management system flowpaths or the electric machine, analyzes the one or more signals to detect a thermal anomaly corresponding to at least one of the electric machine or the thermal management system, and responsive to detecting the thermal anomaly, performs at least one of adjusting a flow of at least one heat exchange fluid of the one or more heat exchange fluids or derating the electric machine.

An aircraft propulsion system includes a gas turbine engine, a nacelle, and an electrical assembly. The gas turbine engine includes a rotational assembly and an inner fixed structure. The nacelle includes a nacelle body, a first bifurcation, and a second bifurcation. The nacelle body extends circumferentially about the gas turbine engine. The nacelle body forms an annular bypass duct between the nacelle body and the inner fixed structure. The first bifurcation and the second bifurcation extend between and connect the nacelle body and the inner fixed structure. The electrical assembly includes a motor-generator, a motor control unit, a motor-generator (MG) cooling system, and a motor control unit (MCU) cooling system. The motor-generator is coupled to the shaft. The motor control unit is electrically connected to the motor-generator. The MG cooling system includes a first heat exchanger disposed at the first bifurcation. The MCU cooling system includes a second heat exchanger disposed at the second bifurcation.

The present invention provides a battery powered fan assembly (1) for providing motive force, comprising: An electric drive motor (12); a fan (14), connected to and driven by the drive motor (12) and rotatable about a central axis X, for moving a fluid F and to thereby create a flow thereof; a battery (16), electrically connected to and for powering the drive motor (12); a heat management system (18), for managing the temperature T of the battery (16) comprising: a fluid collector (22), for collecting collected fluid Fc from the flow of fluid F moved by the fan (14); a heat exchange means (20), for exchanging thermal energy between the collected fluid Fc and the battery (16); and a fluid director (24) for receiving collected fluid Fc from said fluid collector (22) and directing said fluid Fc to said heat exchange means (20) for thermal management of the battery (16).

A VTOL aircraft includes a plurality of tilt propellers configured to be rotated between a forward cruise configuration and a vertical lift configuration. A nacelle or other structure may be configured to form a low pressure zone from the wake of the tilt propellers when operating in the lift configuration. The low pressure zone may include an outlet side of a cooling path. Due to the low pressure, airflow may be induced through the cooling path even in the absence of strong air pressure at the inlet side. During a cruise operation when the VTOL aircraft is in forward motion, the nacelle or other structure may be aligned with a baffle to form an exhaust channel for the cooling airflow.

A VTOL aircraft includes a plurality of tilt propellers configured to rotate between a forward cruise configuration and a vertical lift configuration. A nacelle or other structure may include a variable inlet configured to expand in the lift configuration to maximize the amount of slip flow captured by the inlet. The variable inlet may contract in the cruise configuration to reduce drag. The variable inlet may be passively actuated as the tilt angle of the tilt propeller varies.

The invention discloses an aircraft hybrid electric power cabin thermal management system and a control method thereof, and the aircraft hybrid electric power cabin thermal management system comprises an air incoming flow cooling system, a fuel oil cooling system, an air circulation cooling system, an evaporation circulation cooling system, a power battery heating system and an electric control unit. And the air incoming flow cooling system, the fuel oil cooling system, the air circulation cooling system, the evaporation circulation cooling system and the power battery heating system are in control connection with the electric control unit. According to a current flight stage and an environment temperature state, an air incoming flow cooling system, a fuel oil cooling system, an air circulation cooling system, an evaporation circulation cooling system and a power battery heating system are correspondingly controlled to heat a power battery, a lubricating oil system, a generator, a motor and an electric control unit in a hybrid electric power cabin; and the engine body is subjected to comprehensive thermal management, so that the thermal management requirements of the hybrid electric power cabin in different flight stages and environment states are met, and the hybrid electric power cabin is always in the optimal temperature state.

The invention provides a wide-temperature-range and low-energy-consumption integrated thermal management system for a hovercar. The hovercar integrated heat management system comprises a battery, a motor, an electric controller, an out-cabin radiator, an in-cabin warm air / cold air core body, first to sixth three-way valves, a four-way valve, first to second liquid storage tanks, first to third oil pumps, a gas-liquid separator, a compressor, a condenser, an expansion valve, a plate heat exchanger and the like. Through mutual cooperation and cooperation of the electrically-driven heat dissipation main pipeline, the electrically-driven heat dissipation branch pipeline, the battery heat dissipation main pipeline, the battery heat dissipation first branch pipeline, the battery heat dissipation second branch pipeline and the steam compression circulation pipeline, the heating and heat dissipation requirements of a passenger compartment, the electrically-driven heat dissipation requirement and the battery heat dissipation requirement can be met at the same time, and the system architecture is highly integrated; while the energy utilization efficiency of the whole vehicle is improved, the lightweight level of the whole vehicle is also improved.

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 aircraft propulsion system includes a gas turbine engine, a nacelle, and an electrical assembly. The gas turbine engine includes a rotational assembly and an inner fixed structure. The nacelle includes a nacelle body, a first bifurcation, and a second bifurcation. The nacelle body extends circumferentially about the gas turbine engine. The nacelle body forms an annular bypass duct between the nacelle body and the inner fixed structure. The first bifurcation and the second bifurcation extend between and connect the nacelle body and the inner fixed structure. The electrical assembly includes a motor-generator, a motor control unit, a motor-generator (MG) cooling system, and a motor control unit (MCU) cooling system. The motor-generator is coupled to the shaft. The motor control unit is electrically connected to the motor-generator. The MG cooling system includes a first heat exchanger disposed at the first bifurcation. The MCU cooling system includes a second heat exchanger disposed at the second bifurcation.

The invention provides a propelling device and an aircraft. The propelling device comprises a motor, a movable blade assembly, a stationary blade frame assembly and a duct cylinder with an air duct, and the motor is installed in the air duct through the stationary blade frame assembly; the movable blade assembly is installed on a driving shaft of the motor, and the movable blade assembly is used for generating propulsive force when driven by the motor to rotate and guiding airflow to conduct air cooling heat dissipation on the motor. When the movable blade assemblies rotate, the aircraft can be jointly propelled to run through the pushing force generated by rotation of the movable blade assemblies and the pushing force generated by lift augmentation of the duct cylinders; the air flow entering the air duct can flow through the motor in an accelerated manner in the axial direction of the driving shaft of the motor, so that air-cooled heat dissipation of the motor is realized, and the heat dissipation efficiency of the motor is improved; and the structure is simple, and the weight is reduced.