1031 results about "Cowling" patented technology

An aircraft defining a longitudinal centerline and extending between a forward end and an aft end is provided. The aircraft comprises a fuselage extending between the forward end of the aircraft and the aft end of the aircraft; a wing assembly extending laterally outwardly with respect to the longitudinal centerline from a portion of the fuselage; an unducted turbofan engine including an unducted fan defining a fan diameter; and a pylon fairing that connects the unducted turbofan engine to the wing assembly, the pylon fairing defining a leading edge and including a first inflection point along the leading edge, the first inflection point defining a first radius of curvature that is less than 5 times the fan diameter and greater than 0.1 times the fan diameter.

An aircraft includes a battery pack mounted inside the aircraft, a vent coupled between the battery pack and a surface of the aircraft to at least partly define a vent path between the battery pack and the surface of the aircraft, and a burst membrane located in the vent path. The vent may be coupled to a rear upper portion of a wing or to an outboard side of a nacelle. The aircraft may also include a flexible coupling between the vent and the surface of the aircraft. The aircraft may also include a fairing over a vent outlet to provide a smooth surface for the vent outlet. The battery pack may include battery modules and an enclosure, the battery modules and the enclosure defining paths along which discharge from a thermal event can flow towards the vent.

The invention discloses a liquid cooling heat dissipation ducted fan which comprises a duct, a propeller, a motor, a supporting rib plate, a propeller hub shell and a liquid cooling mechanism. The propeller hub shell is fixed in the duct through the supporting rib plates; a motor and a liquid cooling mechanism are arranged in the propeller hub shell; a rotor output shaft of the motor is connected with a propeller and a fairing through a coupler. The fairing is arranged on the propeller in a sleeving mode, and an opening is formed in the fairing so that blades of the propeller can stretch out. The liquid cooling mechanism is at least partially formed in the propeller hub shell and the supporting rib plate, and the propeller hub shell and the supporting rib plate are of a sandwich structure to form a cooling liquid circulation channel of the liquid cooling mechanism. According to the liquid cooling heat dissipation ducted fan, the cooling liquid flow channels are arranged in the propeller hub shell and the supporting rib plates, the cooling liquid is cooled through the structure of the cooling liquid while the cooling liquid dissipates heat of the motor, the cooling liquid circulates in the propeller hub shell and the supporting rib plates, the motor is continuously cooled, and the heat dissipation efficiency is improved.

The invention discloses a high-heat-dissipation integrated ducted fan electric propulsion system, which belongs to the technical field of ducted fans and comprises a motor controller, a motor, a motor front bearing, a motor rear bearing, a front fairing, a rear fairing and an integrated fan rotating shaft integrally formed by fan blades and a motor rotating shaft. The integrated nacelle motor shell is integrally formed by a motor shell, a ducted nacelle and a plurality of groups of ducted stationary blades; cooling devices are mounted on the motor shell and the motor controller; the cooling device comprises a cooling water channel and a water cooling cavity; a cooling water channel is arranged on the motor shell; the motor controller is provided with a water cooling cavity, and the cooling water channel is communicated with the water cooling cavity; the front fairing is connected to the left end of the integrated fan rotating shaft, and the rear fairing is connected to the right end of the motor controller. The integrated level is high, the size and the weight are reduced by reducing connecting parts, and the power density is high; the system reliability is high; the system cooling efficiency is high; and the power level compatibility and the expandability are high.

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.

A fairing assembly 10 installed about a heat exchanger, such as a radiator R, to reduce the drag on the vehicle, comprising: a frame installed about the heat exchanger, the frame having a curved contour; a panel mounted in the frame, the panel and fairing assembly defining an air passage for air flow to the heat exchanger; a cover installed in the frame and movable to control air flow through the passage to the heat exchanger; and, moving means for moving the cover over the frame to expose a portion of the opening and allow air to flow into the heat exchanger, the frame and cover thereby reducing the amount of drag on the vehicle while still allowing air flow into the heat exchanger; wherein the panel comprises a first overlapping part and a second overlapping part; and wherein the first overlapping part is moveable from an extended position to a retracted position when the fairing is impacted.

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.

A marine drive is for propelling a marine vessel. The marine drive has a propulsor configured to generate a thrust force in a body of water; a battery that powers the propulsor; and a supporting frame which supports the marine drive relative to marine vessel. The supporting frame has a monolithic body defining a frame interior, and further has a support leg extending downwardly from the monolithic body and a steering arm extending forwardly from monolithic body. A cowling is fixed to the supporting frame via at least one hidden fastener that extends from the frame interior, through the supporting frame, and into engagement with the cowl body, wherein hidden fastener being accessible during installation.

The invention discloses a spacecraft fairing environment control system, and belongs to the technical field of ventilation systems. A spacecraft fairing environment control system comprises a gas flow direction control part arranged in a fairing, and further comprises an air environment detection part used for detecting the temperature and humidity in the fairing; the circulating air driving part is used for controlling air flow in the fairing to flow circularly; the air conditioning part is used for humidifying, dehumidifying, heating and cooling the circularly flowing airflow according to the data detected by the air environment detection part; the rapid cooling part is used for discharging low-temperature gas when the temperature of the fairing is rapidly increased; according to the invention, environmental parameters in the fairing can be monitored in real time, and operations such as humidification, dehumidification, heating, cooling and the like can be carried out according to needs, so that the stability of the internal environment of the fairing is kept, and reliable protection is provided for equipment in a spacecraft.

The invention discloses a fairing pulsating pressure scaling wind tunnel test method capable of reducing space-ground difference. The method comprises the following steps: determining initial parameters; based on the determined initial parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test, arranging pulsating pressure measuring points, and carrying out transonic speed state fairing pulsating pressure environment measurement to obtain a measurement result; performing parameter optimization according to a comparison result of a measurement result and a standard parameter under a real flight test condition to obtain an optimal parameter of manual transition; and based on the optimal parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test to obtain test results of the configuration in different free incoming flow Mach number, attack angle and sideslip angle states. According to the method, the free incoming flow local shock wave-boundary layer interference state and the boundary layer separation state of the ground wind tunnel test and the flight test are ensured to be closer through a manual transition mode, and the ground test prediction precision of the carrier rocket fairing pulsating pressure environment is improved.

The invention belongs to the technical field of overall aerodynamic design of helicopters, and particularly relates to an infrared suppressor fairing of a helicopter and an air heat optimization design method of the infrared suppressor fairing. The fairing comprises a fairing body, a shutter air inlet, a square protective net air inlet, a second-stage exhaust pipe and a first-stage mixed exhaust pipe. The front of the fairing body is regarded as a half ellipse which is laterally cut along a long axis; a shutter air inlet is formed in the front portion of the fairing body, and the plane of a shutter is crescent A square protective net air inlet is formed in front of the top of the fairing body; an oval second-stage exhaust pipe is arranged behind the top of the fairing body so that high-temperature exhaust can be further cooled by means of downwash airflow of the rotor wings. A first-stage mixed exhaust pipe is sleeved in the second-stage exhaust pipe; and the primary mixing exhaust pipe is connected with an outlet of the transmitter. According to the method, the problem that the pneumatic performance and the thermal performance are difficult to consider in a traditional design method is solved, comprehensive optimization of the pneumatic performance and the thermal performance of the fairing is achieved, and a pneumatic design scheme of the infrared suppressor fairing is formed.

The invention relates to the technical field of building structures, in particular to a low-carbon energy-saving building structure which comprises a container house, a plurality of ventilation mechanisms used for ventilation are arranged in the container house, each ventilation mechanism comprises a fairing, and a propeller capable of driving air to flow is arranged in each fairing. A drainage mechanism capable of guiding air to enter or be exhausted is arranged at the input end of the fairing, a driving mechanism used for driving the propeller to rotate is fixedly arranged on the upper portion in the container room, and the propeller is arranged at the working end of the driving mechanism. The propeller is driven by the driving mechanism to rotate, so that the propeller drives air to enter the container house through the drainage mechanism, air circulation in the container house is achieved, and the problem that air in a building structure is not circulated is solved.

A propulsion assembly is disclosed including a pylon, a turbomachine having an engine and a fan surrounded by a fan casing, a nacelle surrounding the engine and the fan casing, and including a load bearer positioned at the upper part of the nacelle, a forward fan mount between a frontal zone of the fan casing and a frontal zone of the load bearer, a rear pylon mount between the pylon and the load bearer, and a positioning assembly providing positioning between the fan casing and the load bearer. The propulsion assembly is configured to control the separations between the cowls mounted in a moveable manner on the load bearer and a nose cowl.

The invention discloses an aircraft fairing and a fairing body composite material integrated forming process, and relates to the technical field of aviation equipment manufacturing. The forming process comprises the following steps: 1) preparing a prepreg; (2) sectional forming; 3) a curing process; (4) demolding and trimming; 5) an assembly process; the whole of the front section, the middle section and the tail section of the fairing is of a long-size U-shaped streamline structure. According to the invention, through molecular collaborative design of fluorination modification-polysiloxane toughening-cyanate ester interpenetration, the key contradiction of a high-performance fairing material is systematically solved; ultrahigh strength and toughness are realized at the same time by utilizing in-situ nano microphase separation; a dense graphite carbon layer is generated through catalysis, so that efficient halogen-free flame retardance is achieved; by means of the partition design of molecular fluorination and surface metallization, the high wave transmission and lightning protection functions are ingeniously unified. And finally, a new generation of aircraft fairing composite material with force bearing, wave transmission, heat resistance, impact resistance and environmental stability is successfully prepared.

The embodiment of the invention discloses a digitalized assembling method of a helicopter fairing, which comprises the following steps of: 1, scanning the surface of the fairing and the surface of a helicopter body structure, and acquiring a three-dimensional contour model of the surface of the fairing and the surface of the helicopter body structure according to scanning data; 2, generating a visual assembly correction scheme by comparing the three-dimensional contour model with a design model imported into a controller; 3, in the fairing assembly process, a visual assembly correction scheme is projected through laser, non-contact hole making and polishing guiding are achieved, and digital assembly of the fairing is implemented; and 4, in the fairing assembling process, dynamic position information of the fairing and all the positioning targets on the aircraft body structure is obtained, the moving track of the fairing in the assembling process is solved, automatic alarming is conducted when the fairing is out of tolerance, and laser projection is frozen. The problems that an existing fairing assembling mode is low in assembling efficiency and precision, and the production efficiency and the assembling requirement of a helicopter are difficult to meet are solved.

The invention relates to the field of aero-engines, and discloses an anti-icing-surge valve grid device for an aero-engine and an adjusting method. The device comprises a shell, a valve, a fairing, a connecting shell, a piston rod and a connecting base, grids with different numbers and different areas are arranged in the shell in the circumferential direction, and a first anti-icing outlet, an air outlet, an air inlet and a second anti-icing outlet are sequentially formed in the outer wall of the shell; the piston rod is used for driving the valve and the fairing which are connected with the piston rod to move in the shell, so that the elliptical hole in the valve, the curved hole in the fairing and the grids with different areas work cooperatively to form different opening degrees, airflow entering from the air inlet passes through the different opening degrees, and the air flow is formed. Therefore, the flow flowing into the first anti-icing outlet, the second anti-icing outlet and the air outlet is controlled, switching of the anti-icing function and the anti-icing surge function and flow adaptation are achieved, and the energy utilization and reliability guarantee system of an air path system in an aero-engine is optimized.

A firewall system isolating a fire zone of an engine (2) from a hanger (1) located in a non-fire zone, comprising: a vertical firewall (3) arranged at a front end of a frame of the hanger, on which a support structure (13) is arranged for mounting the hanger and a fairing of the engine; and a bottom firewall (7) which is provided so as to extend rearward from the vertical firewall, is provided so as to cover the suspended frame from below, has a bottom plate (71), a left side plate (72), and a right side plate (73), and has an isolated space formed by the vertical firewall, the bottom firewall, and the suspended frame. Fireproof sealing isolation of an aircraft engine can be achieved, meanwhile, arrangement of multiple systems such as fuel oil, hydraulic pressure, an air source, fire prevention, power and EWIS is achieved, influences between the systems are avoided, and structural weight reduction can be achieved.

The invention provides an unmanned aerial vehicle propeller power and acoustic performance test system, and the system comprises a vertical column, the bottom of which is fixed on the ground through a supporting structure; the first cover closing type switching disc is arranged at the top end of the stand column; the six-axis torque sensor is arranged at the top end of the first cover closing type transfer disc; the second closed cover type adapter disc is arranged at the top end of the six-axis torque sensor, and an accommodating cavity is formed in the second closed cover type adapter disc; the magnetic encoder is fixedly mounted in the accommodating cavity; the motor is arranged at the top end of the second cover closing type transfer disc; the propeller is arranged on the motor shaft in a sleeving manner; the fairing is arranged outside the stand column in a covering mode; the electronic speed regulator is used for controlling the motor to rotate when the starting instruction is received; and the analysis module is used for acquiring force data and torque data output by the six-axis torque sensor and sound pressure data detected by the microphone, and obtaining a time domain graph, a frequency domain graph, a data table and a curve. The beneficial effects of the invention are that the device can more accurately measure and analyze the force, the torque and the radiated sound pressure of the propeller.

The invention discloses a combustion chamber air fairing which comprises a fairing outer cylinder (1) and a guide plate (2), the guide plate (2) is located on the radial inner side of the fairing outer cylinder (1), at least part of the guide plate (2) is a curved surface, a plurality of through holes are formed in the fairing outer cylinder (1), and the combustion chamber air fairing is characterized by further comprising a plurality of partition plates (3) connected to the fairing outer cylinder (1) and the guide plate (2) and further comprising spoilers (4). The problems that uniformity of air flowing into different nozzles from a fairing is poor, a thin-wall structure of the fairing is prone to being damaged during operation, and the mixing effect of air and fuel in a combustion chamber is common are solved, meanwhile, vortex and turbulent flow of the air before the air enters the nozzles are reduced, air flowing is more stable, energy loss in the combustion chamber is reduced, combustion efficiency is improved, and the service life of the combustion chamber is prolonged. The stability and the uniformity of the combustion process are ensured; the invention further discloses a combustion chamber with the combustion chamber air fairing.

The invention discloses a wind turbine fairing optimization design method, which relates to the technical field of fluid machinery and engineering, and comprises the following steps: a basic parameter acquisition and working condition defining step: collecting wind turbine model parameters, and measuring the wind condition of an operation area; the flow guide cover is divided into a windward section, a middle section and a wake flow section; a multi-objective optimization system construction step: determining aerodynamic resistance, structural strength, weight and wake flow influence optimization objectives; an optimization algorithm iteration solving step, wherein an improved non-dominated sorting genetic algorithm is adopted; a simulation verification and parameter correction step: selecting a scheme to perform CFD and finite element simulation; and a physical prototype manufacturing and performance testing step: manufacturing a prototype, and carrying out wind tunnel and field testing. The design efficiency of the air guide sleeve is improved, systematicness is optimized, pneumatic, structure and weight multi-target balance is achieved, cross-model universality is improved, and the problems that traditional design is dependent on experience and unbalanced in performance are solved.

A solid carrier fairing that can resist pulsating pressure under large cone angles includes: an end, a front cone section, a column section, an inverted cone section, a segmented longitudinal separation structure, a combined exhaust structure, and a bolt box docking frame combined structure. The end, front cone section, column section, and inverted cone section are all column structures formed by combining two half-covers, which are connected to the segmented longitudinal separation structure internally, and the end, front cone section, column section, and inverted cone section are connected in sequence through end faces to form a solid carrier fairing structure. The segmented longitudinal separation structure is used to make the solid carrier fairing structure complete before separation, and after separation, the solid carrier fairing is changed into two half-covers to ensure normal separation of the fairing; a combined exhaust structure and a bolt box docking frame combined structure are provided at the tail end of the inverted cone section to avoid the exhaust structure from affecting the internal noise. The combined exhaust structure is used to balance the pressure difference between the inside and outside of the fairing during the flight of the solid carrier rocket, and the bolt box docking frame combined structure is used to connect and carry the fairing and the control cabin.

Aircraft comprise a fuselage, wings coupled to the fuselage, a wing-to-body fairing operatively coupled relative to the fuselage and the wings and defining a fairing volume between the wing-to-body fairing and at least the fuselage, and a fairing fuel tank positioned within the fairing volume.

Propeller restraint kits for turboprop aircraft are provided which include at least one, and preferably a pair, of propeller restraint systems. Each propeller restraint system will preferably include an open-ended propeller restraint boot that may be formed of a flexible and durable fabric material which is sized and configured for placement over the tip of a propeller blade. A cowling attachment assembly is provided in the kit and is adapted to being attached to an aircraft cowling skin. A propeller boot tether strap of predetermined length interconnects the propeller restraint boot and the cowling attachment assembly. The cowling attachment assembly includes (i) a quick disconnect (QD) pin assembly connected to a terminal end of the propeller boot tether strap, and (ii) a pin receiver adapted to being positionally attached to an interior portion of the aircraft cowling skin.

The invention belongs to the technical field of aircraft cockpit windshield glass structure design, and particularly relates to an aircraft cockpit windshield glass structure which comprises an electromagnetic film, outer-layer glass, inner-layer glass, a metal frame, a metal sinking frame and a curved-surface wave-transparent fairing. The electromagnetic film is positioned between the outer-layer glass and the inner-layer glass; the inner-layer glass is arranged in the metal frame; the electromagnetic film is in flush contact with the outer side surface of the metal frame; the metal sinking frame is mounted in the windshield window, and the cross section of the metal sinking frame is L-shaped; the metal frame is arranged in the L-shaped opening of the metal sinking frame, the outer side surfaces of the metal frame and the metal sinking frame are flush and contacted with the aircraft body surface skin, and the outer-layer glass protrudes out of the aircraft body surface skin; the curved-surface wave-transparent fairing surrounds the periphery of the outer-layer glass and is in smooth transition between the outer-layer glass and the surface skin of the machine body.

The invention discloses a built-in fairing pre-tightening and rotary separation hinge mechanism, relates to the technical field of carrier rockets, and solves the problems that a fairing separation mode is complex in structure, an exposed separation mechanism affects the aerodynamic configuration of an aircraft, and an axial pre-tightening function does not exist. An upper fairing is movably connected with a cabin body; a hinge structure is arranged on the upper fairing, and a cabin body groove structure is arranged on the cabin body; one end of the hinge structure is inserted into an insertion groove in the cabin groove structure, the other end of the hinge structure is provided with a pre-tightening structure, the pre-tightening structure abuts against the inner wall of the cabin groove structure, and the hinge structure, the pre-tightening structure and the cabin groove are matched to achieve circumferential positioning and axial pre-tightening of the upper fairing; the cabin body is provided with a step structure, and the upper fairing is buckled on the step structure. According to the device, the circumferential positioning and axial pre-tightening functions of the fairing can be achieved, the hinge is used as a turnover fulcrum to ensure that the turnover track is controllable, and the situation of collision or dead locking is avoided; and meanwhile, the influence of the separation mechanism on the aerodynamic configuration of the aircraft is reduced.

The invention provides an intelligent simulation training bomb for ground attack of an unmanned aerial vehicle. The training bomb comprises a bomb body, a thrust reverser, a fairing, a mounting frame, a smoke tank and a control module. Wherein a gas cabin is arranged in the bomb body, and the thrust reverser is composed of an electromagnetic valve and a multi-way pipe. The mounting frame integrates a fuming tank, a control module and a motion sensing module, and the control module calculates the falling speed, the grounding time and the position coordinates in real time based on data of all sensors of the motion sensing module so as to control the electromagnetic valve to be opened, so that compressed gas in the gas cabin is ejected from a reverse thrust gas opening in the circumferential direction of the fairing through the multi-way pipe; in combination with deformation buffering of the fairing made of an elastic material, nondestructive recovery of the training bomb is achieved, and repeated use is facilitated. And meanwhile, the fuming tank is controlled to trigger and release tracing smoke, so that an impact point is marked, and parameters such as latitude and longitude during landing are sent to a ground guide training system to serve as an evaluation basis of throwing precision.

The invention belongs to the field of unmanned aerial vehicle structures, and discloses an unmanned aerial vehicle and a fuselage fairing structure thereof, the unmanned aerial vehicle comprises a cover body, the top of the cover body is provided with a hinge joint connected with a fuselage of the unmanned aerial vehicle, the cover body is also provided with a quick release lock connected with the fuselage of the unmanned aerial vehicle, and the cover body is internally provided with an air support rod connected with the fuselage of the unmanned aerial vehicle; when the quick-release lock releases the connection between the cover body and the fuselage of the unmanned aerial vehicle, the air supporting rod can upwards support the cover body, and when the air supporting rod upwards supports the cover body, the cover body rotates around the hinge joint. The fuselage fairing structure of the unmanned aerial vehicle is convenient to open, and then maintenance of the large freight unmanned aerial vehicle is facilitated.

A combustor for a gas turbine engine includes a ceramic matrix composite (CMC) dome structure, a CMC outer liner, a CMC inner liner, and a cowl structure, that are connected together via at least one outer connection and at least one inner connection. An outer seal member is arranged to seal an outer gap between an outer liner connecting flange and an outer dome connecting flange, and an inner seal member is arranged to seal an inner gap between an inner liner connecting flange and an inner dome connecting flange.

The high-speed engine blade deformation measurement system and measurement method for wind tunnel test belong to the technical field of full-surface deformation measurement of engine blades. In order to realize the full-surface deformation measurement of blades, the present invention includes a first imaging device and a second imaging device connected to a digital pulse signal delay generator and a computer respectively, the digital pulse signal delay generator is connected to a laser, the laser is connected to a beam expander, the beam expander is connected to an optical fiber image transmission bundle, and the optical fiber image transmission bundle is connected to a zoom lens; the beam expander, the optical fiber image transmission bundle, and the zoom lens constitute a laser homogenization device, and the laser emitted by the zoom lens is irradiated onto the blades and the fairing; the first imaging device and the second imaging device are aligned with the position where the laser irradiates the high-speed engine blades to shoot. The present invention significantly improves the measurement accuracy and the comprehensiveness of the data, and improves the overall performance and reliability of the engine.