2116 results about "Nacelle" patented technology

A wind energy plant with a nacelle, a rotor, which features at least one rotor blade adjustable in its blade pitch angle, a central control device for controlling the wind energy plant and a control unit disposed in the rotor for controlling the blade pitch angle of the at least one rotor blade, wherein the central control device and the control unit in the rotor can exchange data with each other via a data link, which comprises at least one first sending and receiving device at the nacelle side and at least one second sending and receiving device at the rotor side, wherein a wireless network connection with a safety-oriented communication protocol is provided between the at least one first sending and receiving device and the at least one second sending and receiving device, and that one monitoring- and communication device at a time is associated to the first and / or the second sending and receiving device, which can monitor the function of the sending and receiving device and initiate a predetermined action in the case of an error.

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.

The invention belongs to the technical field of engine nacelle part production, and particularly relates to an intelligent welding system and method for an engine nacelle combined type air inlet fairing. An inner spoiler moving groove and an outer spoiler moving groove are machined in the air inlet duct casting blank borne by the air inlet duct positioning assembly through the inner cutting assembly and the outer cutting assembly correspondingly, and the air inlet duct casting blank is formed into an air inlet duct component; a first clamping feeding assembly, a second clamping feeding assembly, a third clamping feeding assembly and a fourth clamping feeding assembly are used for feeding an inner reinforcing part, a heating part, a noise reduction and sound absorption part preassembled with a spoiler assembly and an annular cover plate into a cavity of the air inlet channel component, and the fed parts are welded and fixed into the cavity of the air inlet channel component through a welding assembly; in this way, intelligent welding and assembling operation of the combined air inlet fairing can be achieved, and the combined air inlet fairing has the advantages of being high in structural strength, good in noise reduction and sound absorption effect and capable of achieving automatic deicing of the lip and air inlet distortion control.

The utility model discloses an engine nacelle, which is applied to the field of aircraft manufacturing, is used for improving the structural strength of an aircraft engine nacelle, and comprises a front firewall, a rear firewall, an upper firewall, a front section skin, a middle section skin, a rear section skin and a mounting frame, a second through hole is formed in the rear firewall; the upper firewall is arranged between the front firewall and the rear firewall, the front end of the upper firewall is connected with the front firewall, the rear end of the upper firewall is connected with the rear firewall, the mounting frame and the U-shaped reinforcing beam are both of a U-shaped structure and are both located on the front side of the front firewall, and the mounting frame is arranged above the U-shaped reinforcing beam; the rear end of a front-section lower wall plate of the front-section skin is connected with the U-shaped reinforcing beam; the front ends of a middle-section left wall plate, a middle-section right wall plate and a middle-section upper wall plate of the middle-section skin are connected with the mounting frame; and the outer sides of the upper firewall, the front firewall and the rear firewall are coated with the middle-section skin.

The invention discloses a floating type offshore wind turbine aerodynamic load calculation method based on a frequency domain method. The method comprises the steps that a state derivative is determined based on a blade element momentum theory quasi-steady-state assumption, and a dynamic inflow model is used for correction; establishing a pneumatic-elastic coupling two-degree-of-freedom linear dynamic model; establishing a wind wheel kinetic model; establishing a control model according to a floating type offshore wind turbine pitch angle control strategy, a torque control strategy and a floating body motion feedback control strategy; based on the pneumatic-elastic coupling two-degree-of-freedom linear dynamic model, the wind wheel dynamic model and the control model, pneumatic additional mass and pneumatic additional damping are obtained; pneumatic exciting force acting on the hub when the position of the cabin is fixed is obtained; and based on the pneumatic additional mass, the pneumatic additional damping and the pneumatic exciting force, an overall linear dynamic model of the fan is established and solved, and overall response is obtained. According to the method, the calculation result precision can be improved while the calculation efficiency is ensured, and a technical support is provided for accelerating the simulation efficiency of the offshore wind turbine.

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 invention belongs to the technical field of engine nacelles, and particularly relates to an engine nacelle fairing with an inner supporting framework system.The engine nacelle fairing is reasonable in design and is mainly composed of a rectifying outer cover, a first inner supporting framework, a second inner supporting framework, an arc-shaped air guide pipe and an annular cover plate; a plurality of sets of first positioning protrusions and second positioning protrusions corresponding to the first positioning protrusions are installed in an annular cavity of the rectification outer cover, and the first inner supporting frameworks or the second inner supporting frameworks are connected in an inserted mode through the first positioning protrusions and the second positioning protrusions corresponding to the first positioning protrusions. The first inner supporting framework, the second inner supporting framework and the arc-shaped air guide pipe can be pre-assembled into an inner supporting framework system, the annular cover plate is used for sealing the annular cavity of the rectification outer cover, in this way, the overall compressive strength of the engine nacelle fairing can be greatly improved, and meanwhile heat generated by working of an engine can be used for achieving freezing prevention of the fairing.

The invention belongs to the field of flight control, and particularly relates to a tilt-rotor aircraft transition state undisturbed switching control method based on linear active disturbance rejection. Longitudinal dynamics of the tilt rotorcraft is modeled into a switching system, a nacelle inclination angle is combined to serve as a switching signal, a switching active-disturbance-rejection controller is designed, internal / external disturbance of the system is estimated in real time and compensated, and multi-channel decoupling is achieved; and a variable gain continuous switching controller is further introduced, and a time interval is set and adjusted after a switching moment, so that jolt at the switching moment is inhibited, and the influence of controller switching on the working process of the conversion process of the flight mode of the tilt-rotor aircraft is effectively reduced. Based on a dynamic error system and a multi-Lyapunov function, in combination with an average residence time constraint, a linear matrix inequality condition for ensuring the asymptotic stability of the system is deduced, and a simulation experiment of the transition state of the tilt-rotor aircraft is carried out.

The invention provides an aviation nacelle part waviness automatic detection and evaluation method, and belongs to the field of product production detection, and the method comprises the steps: carrying out the protocol scanning of part surface contour features, and obtaining the undistorted coordinate point discrete point cloud data of the actual surface contour; filtering the point cloud data; obtaining a piece of flattened deviation value discrete point cloud data; obtaining a flattened deviation value continuous curve by adopting a general standard curve reduction method; extracting waviness components, and obtaining a curve center line; obtaining a crest value and a trough value participating in calculation; and calculating the maximum wave amplitude, the waviness and the slope for evaluating the waviness by using the wave peak value and the wave trough value. According to the method, the relation between the minimum cut-off wavelength of the waviness and the sampling interval is stipulated, the undistorted reduction of the surface profile curve is ensured, the influence of non-waviness profile components in the digital evaluation of the waviness is eliminated through a combined calculation method, and the evaluation accuracy of the waviness of the aviation nacelle part is improved.

An off-grid wind turbine system comprising a wind turbine with an electric generator for generating an initial electric power output. An electrolyzer system with a hydrogen electrolyzer located inside the nacelle or tower of the wind turbine, so as to generate hydrogen by an electrolysis process. An electric converter system serves to convert the initial electric power output into a DC electric power output dedicated for powering the electrolyzer. The produced hydrogen is stored in a hydrogen storage tank.

The invention discloses an impeller unbalance monitoring method based on a blade three-point identification visual technology. The impeller unbalance monitoring method comprises the steps that three identification points are marked on each blade; in the working process of the wind turbine generator, the cabin clearance monitoring camera is used for shooting images of the blades in real time; identifying the pixel coordinates of the identification points in the image through an image technology, and converting the pixel coordinates into world coordinates; fitting a blade track according to the world coordinates; and identifying the impeller imbalance condition according to the blade motion trail. According to the method, image data collected by a cabin clearance monitoring camera (located at the bottom of a cabin and close to a hub shooting blade) is utilized, pixel coordinates of marking points and internal and external parameters of the camera are recognized through an image technology, the positions of the marking points in a world coordinate system are converted, and the position of the marking points is determined according to the difference of tracks of the marking points of the three blades. And identifying the impeller imbalance condition of the wind generating set.

The present application discloses an aircraft configured for vertical take-off and landing and horizontal flight. The aircraft includes (i) a first nacelle fuel tank configured to hold a first fuel supply for a first engine and to selectively provide fuel to the first engine during both (a) a vertical take-off and landing, and (b) horizontal flight; (ii) a fuselage fuel tank, and (iii) a control system configured to redistribute fuel between the first nacelle fuel tank and the fuselage fuel tank.

The invention relates to the technical field of intelligent unmanned ships, in particular to a wave self-adaptive unmanned ship, a self-adaptive control system and a control method.The wave self-adaptive unmanned ship comprises a main ship body, a front arch, a rear arch, a suspension system, a propelling nacelle, a rigid sliding plate and a buoy, the main ship body is connected with the rear arch and the front arch, and the rigid sliding plate is fixed to the buoy; the propulsion nacelle is installed at the tail of the rigid sliding plate, the rear arch and the rigid sliding plate are installed through two bearings perpendicular to each other, the suspension system comprises a rocker arm, a plate spring and a TPMLM actuator, the rocker arm is connected to the rigid sliding plate, the plate spring and the TPMLM actuator are connected between the rocker arm and the rigid sliding plate and support the rocker arm to be in an upward inclined posture, and the rear arch is connected to the rigid sliding plate. The front arch is connected with the rocker arm, the plate spring and the TPMLM actuator are installed in parallel in the vertical direction, the TPMLM actuator is used for outputting active control force for restraining movement of the ship body, and the unmanned ship, the self-adaptive control system and the self-adaptive control method have the advantages that wave loads can be actively counteracted, and therefore movement restraining is achieved.

An arrangement for fixing a housing of a wind turbine gearbox in a nacelle, including a housing-fixing means, a nacelle-fixing means, and at least two spring devices. The spring devices support the housing-fixing means against the nacelle-fixing means in opposite directions along a support axis. The spring devices are arranged on different sides of a plane extending parallel to an axis of rotation of an input shaft of the wind turbine gearbox.

The invention discloses a yaw wind error compensation method and system based on wind measurement data, and relates to the technical field of wind power generation equipment control, and the method comprises the steps: setting a wind measurement point position, and collecting an original wind field data set; calculating an incoming flow and a cabin wind direction angle according to the original wind field data set; performing deviation calculation to obtain a yaw wind error angle; triggering a dual-channel compensation module, accumulating errors, constructing an error queue, and performing batch compensation to obtain a compensation result; yawing control parameters are called, the cabin attitude is analyzed according to the compensation result, the yawing control parameters are updated, and self-adaptive compensation of yawing to the wind error is achieved. The technical problems that in the operation process of the wind driven generator, due to wake flow, terrain, turbulent flow and other environmental factors and equipment loss, a traditional yaw system cannot accurately detect the wind direction and compensate the wind alignment error, the power generation efficiency is low, and the potential safety hazard of equipment is large are solved, and the purposes of accurately improving the wind alignment precision of a wind wheel and improving the power generation efficiency are achieved. The power generation efficiency is obviously improved; and the equipment operation safety is enhanced.

The invention discloses a flying wing solar unmanned aerial vehicle based on distributed reversible power omni-directional cooperative control, and relates to the technical field of aircraft design, the flying wing solar unmanned aerial vehicle comprises a cabin, flying wing platforms are symmetrically arranged on the two sides of the cabin, reversible power assemblies are arranged on the flying wing platforms, and each reversible power assembly comprises a supporting unit and a driving unit; a multi-mode flight control assembly is arranged in the cabin. According to the flying wing solar unmanned aerial vehicle based on distributed reversible power omni-directional cooperative control, in the air flight stage, based on the flexible characteristic of the flying wing platform, the driving units are vertically arranged in pairs, direct force is used for driving, the characteristic of multi-mode control is embodied, and the reversible power assembly is used for cooperatively achieving the omni-directional operation function; in the ground take-off and landing stage, based on the large deformation characteristic of the flying wing platform, the space distribution form of the driving units is adjusted in a reversing mode through the supporting units, every two driving units are horizontally arranged, control is achieved through wing deformation, and therefore the power units are prevented from colliding with the ground.

The invention discloses a rotorcraft flutter model design and test method for a low-speed wind tunnel test, and belongs to the technical field of low-speed wind tunnel tests. The method solves the difficult problem of how to separate key nacelle modes in rotation flutter of the tilt rotorcraft. A fuselage is connected with the side wall of a wind tunnel, wings are of a double-girder straight wing structure, one end of each wing is connected with the fuselage, the other end of each wing is connected with a nacelle mechanism, a rotating paddle is connected with the nacelle mechanism, and a motor, a yawing mechanism and a pitching mechanism are arranged in the nacelle mechanism. The yawing mechanism and the pitching mechanism are connected in a matched mode to achieve decoupling in the yawing direction and the pitching direction of the model, the motor is installed on the pitching mechanism, and the execution end of the motor is connected with a main shaft of the rotating paddle. According to the rotorcraft flutter model design and test method for the low-speed wind tunnel test, decoupling of modes in the pitching direction and the yawing direction of the model can be achieved, the flutter characteristic of the model is shown as coupling of the pitching mode and the yawing mode of a nacelle, and the test requirement of the model wind tunnel test is met.

The invention relates to the technical field of wind turbines, and particularly provides a sliding bearing system and a maintenance method thereof.The sliding bearing system comprises a pair of radial sliding bearings located at the hub end and the gearbox end of a main shaft respectively, and the main shaft is rotatably connected to a cabin through the pair of radial sliding bearings; the axial sliding bearing is located on the opposite outer side of the gear box end of the main shaft, and the axial sliding bearing is connected with the cabin and the main shaft; the axial sliding bearing includes an annular plate detachably connected to the gearbox end of the main shaft and a number of caliper slider modules arranged spaced apart from each other along a circumferential direction of the annular plate and detachably connected to the nacelle and each releasably clamping the annular plate. Through the configuration, the caliper type sliding block module can loosen and clamp the annular plate, so that the caliper type sliding block module can be loosened and further pulled out from the annular plate, and the axial sliding bearing with the replaceable sliding blocks on the two sides is realized.

The invention belongs to the technical field of unmanned aerial vehicles, and discloses a foldable multi-rotor hybrid power unmanned aerial vehicle which comprises lift rotors (10), attitude rotors (20), vehicle arms (30), a cabin (40), a photoelectric pod (50), a cargo hold (60) and an undercarriage (70). The lift rotors and the attitude rotors are combined, the attitude control difficulty is reduced while the large load is guaranteed, and compared with a traditional helicopter, the mechanical structure complexity is reduced, and maintenance and replacement are convenient; and when the power system is redundant, the reliability of the method is ensured through a system fault reconstruction scheme.

The invention discloses an energy storage balance wind power generation control system, and belongs to the technical field of energy storage and power supply, the energy storage balance wind power generation control system is provided with an energy storage adjusting structure and a speed reduction structure, the energy storage adjusting structure comprises a transformer T, and an adjustable slide rheostat is arranged between each phase input end of the transformer T and each phase output end of an alternating current generator; by automatically adjusting the resistance value of the slide rheostat, the voltage of each phase input end of the transformer T is automatically stabilized. The speed reduction structure comprises a cabin, a rotating shaft is arranged in the cabin, one end of the rotating shaft is fixedly connected with blades, the other end of the rotating shaft is fixedly connected with an alternating current generator, a rotating disc is fixedly connected to the surface of the rotating shaft, a rotating disc groove is formed in the rotating disc, and a weight block is arranged in the rotating disc groove. The arc-shaped block is arranged on the surface of the weight block and abuts against the inner wall of the cabin, so that the rotating speed of the rotating shaft and the blades is reduced, the voltage generated by the alternating-current generator is more stable, and the voltage balance when wind power generation is merged into a municipal power grid is improved.

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.

This invention pertains to aircraft seating, specifically a double-decker semi-reclining airplane seat designed to enhance passenger comfort in economy class. It comprises a front support frame, a rear support frame, a lower seat, and an upper seat. The support frames are positioned parallel and obliquely on a slide rail. The lower seat, situated between the frames, features a movable seatback and a hinged seat base. The upper seat includes a hinged seatback, seat base, and leg support, with a footrest platform hinged to the front support frame. This design optimizes limited cabin space while providing a more ergonomic seating arrangement. Evenly distributing body weight reduces pressure on the waist, hips, and legs, alleviating discomfort from prolonged sitting. Compared to traditional economy seats, this innovation offers improved reclining support, enhancing overall passenger comfort during long flights.

The invention discloses a reusable anti-unmanned aerial vehicle aircraft based on rotary cutting steel blades, and belongs to the technical field of unmanned aerial vehicle countering. The aircraft comprises a seeker cabin, a function cabin, an instrument cabin, a power cabin, a steering engine cabin and a parachute pack cabin. An infrared seeker is arranged in the seeker cabin; the functional cabin comprises a fixed cabin body, a rotating cabin body and a folding and unfolding mechanism, a fixed shaft is arranged at the axis position of the fixed cabin body, N folding steel blades are installed on the rotating cabin body and evenly distributed in the circumferential direction, and the folding and unfolding mechanism drives the folding steel blades to complete unfolding and folding actions. The folding and unfolding mechanism simultaneously drives the rotary cabin to rotate around the axis of the aircraft; a main power supply, an airborne computer, a data chain and an inertial navigation device are arranged in the instrument cabin, and the power cabin provides patrol flight power for the aircraft; a steering engine for controlling the rudder wings to rotate is arranged in the cabin, and a pressurizing device for pressurizing the parachute pack cabin is arranged in the parachute pack cabin. The anti-unmanned aerial vehicle aircraft can be repeatedly used, and the combat time of the aircraft is prolonged to the maximum extent.

In a first aspect of the present invention there is provided a wind turbine comprising a tower, a nacelle mounted on the tower, and a rotor mounted to the nacelle. The rotor comprises a hub and at least three wind turbine blades. Each blade extends between a root and a tip. Each blade further comprises a connection point located between the root and the tip. The wind turbine further comprises a plurality of blade connecting members, each blade connecting member being connected between corresponding connection points of a pair of wind turbine blades. The wind turbine further comprises a tensioning system for adjusting the tension in each blade connecting member. The tensioning system comprises a plurality of linear actuators, each linear actuator being coupled between the hub and a respective blade connecting member. Each linear actuator is configured to adjust the tension in the blade connecting member.

The invention discloses an anti-unmanned aerial vehicle aircraft based on folding steel blade rotary cutting, and belongs to the technical field of unmanned aerial vehicle countering. The aircraft comprises a seeker cabin, a function cabin, an instrument cabin, a power cabin, a steering engine cabin and a parachute pack cabin. An infrared seeker is arranged in the seeker cabin; the functional cabin can independently rotate relative to the front cabin section and the rear cabin section under the driving of the motor and release the folding steel blade mounted on the surface of the functional cabin in a folding manner, and the motor drives the functional cabin to rotate at a high speed and drives the folding steel blade to rotate, so that rotary cutting of the target unmanned aerial vehicle is realized; a main power supply, an airborne computer, a data chain and an inertial navigation device are arranged in the instrument cabin, and the power cabin provides patrol flight power for the aircraft; a steering engine for controlling the rudder wings to rotate is arranged in the cabin, and a pressurizing device for pressurizing the parachute pack cabin is arranged in the parachute pack cabin. The anti-unmanned aerial vehicle aircraft is few in component parts, high in reliability, low in cost and high in cost-effectiveness ratio.

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 present invention provides an aviation pilot seat with a lightweight and high-load side-slip adjustment mechanism, the seat comprising a leg assembly and a guide rail assembly, a seat assembly mounted on the leg assembly, the seat assembly comprising a seat cushion and a side-slip adjustment mechanism arranged under the seat cushion. The left and right support plates of the seat are rationally split, optimized by simulation and verified by tests, which can not only provide a safe and reliable side-slip adjustment function for the aircraft pilot, but also maintain the force and strength requirements of the seat. The integrated design cleverly avoids the cantilever load condition before side-slipping, achieves the optimal force structure, and maximizes the lightweight and high-load design goal of the seat, without affecting the verification of various load tests of the seat. Even during a crash impact, the seat can still firmly restrain the pilot; in order to prevent the pilot from operating incorrectly, a front-back adjustment limiting mechanism is added after the side-slip adjustment to prevent the front-back adjustment from hitting the cabin side wall after the side-slip adjustment.

The invention belongs to the field of vertical take-off and landing aircrafts, and relates to a distributed contra-rotating rotor hybrid tilting rotorcraft. Comprising a front tilting electric contra-rotating rotor nacelle, a front cabin body, a wing end tilting electric contra-rotating rotor nacelle, a V-shaped empennage, a V-shaped empennage aileron, a vertical empennage, a retractable undercarriage, a green hybrid power energy system and an empennage tilting electric contra-rotating rotor nacelle, the front tilting electric contra-rotating rotor nacelle is hinged with the other end of the front cabin body; an empennage tilting electric contra-rotating rotor nacelle is arranged at the end part of the V-shaped empennage; wing end tilting electric contra-rotating rotor wing nacelles are arranged at the ends of the wings; and rotor wing units are arranged in the rotor wing nacelles, the six rotor wing units all adopt a coaxial reverse upper and lower two-paddle structure, and the rotating directions of the rotor wings in the six rotor wing units are that the adjacent rotating directions and the opposite rotating directions are opposite.

The invention discloses an anti-unmanned aerial vehicle aircraft based on rotary cutting, and belongs to the technical field of unmanned aerial vehicle countering. The aircraft comprises a seeker cabin, a steel rope cabin, an instrument cabin, a power cabin, a steering engine cabin and a parachute pack cabin. An infrared seeker is arranged in the seeker cabin; the steel rope cabin can independently rotate relative to the front cabin section and the rear cabin section under driving of the motor and release the flexible steel rope wound in the steel rope cabin, the motor drives a shell of the steel rope cabin to rotate at a high speed and drives the flexible steel rope to rotate at a high speed, and therefore rotary cutting of the target unmanned aerial vehicle is achieved. A main power supply, a control computer, a data chain and an inertial navigation device are arranged in the instrument cabin; the power cabin provides patrol flight power for the aircraft; a steering engine for controlling the rudder wings to rotate is arranged in the cabin, and a pressurizing device for pressurizing the parachute pack cabin is arranged in the parachute pack cabin; the target unmanned aerial vehicle is softly killed, the application range is wide, and the problem that the cost-effectiveness ratio of an existing anti-unmanned aerial vehicle aircraft is relatively low can be effectively solved.

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.