62results about How to "Guaranteed aerodynamic performance" patented technology

The invention relates to a draw bail of a vane wheel and a high-speed rotation shaft of a high-speed centrifugal compression device. The invention ensures that the high-speed rotation shaft and the vane wheel are not deformed and the assembly and disassembly are convenient. The draw bail comprises the high-speed rotation shaft and the vane wheel and is characterized in that a stretching shaft penetrates through a centre hole of the vane wheel, one end of the stretching shaft is connected with the inner screwed hole on the terminal of the high-speed rotation shaft by screw thread, a gland nut is sheathed on the other end of the stretching shaft, the end surface is fixedly connected with an outrigger shaft neck of the vane wheel, and the end surface of one end of the gland nut is fixedly connected with the end surface of the vane wheel.

The invention discloses an energy absorption structure for a corrugated sandwich panel. The energy absorption structure consists of an upper panel, multiple layers of corrugated cores, a sandwich panel and a lower panel, all of which are sequentially stacked and are adhered by an epoxy-structure adhesive or brazed to form an integrated structure; the section of the corrugated core is in a trapezoid, triangle or sine curve shape, and when the layer number of the corrugated cores is larger than or equal to two, the arrangement manner of the corrugated cores is regular arrangement, staggered arrangement or 0-degree or 90-degree arrangement. The energy absorption structure for the corrugated sandwich panel has the characteristics of light weight, strong toughness and good shock resistance; in the whole energy absorption procedure, the impact force is stable; due to the openness of a core material structure, the energy absorption structure can effectively and actively dissipate collision heat during collision, so that the structure material avoids yielding in advance under the influence of high temperature. The energy absorption structure for the corrugated sandwich panel can be applied to an automobile buffering energy absorption device, and also can be applied to any load and field needing the buffering energy absorption structure.

The invention provides a U-shaped wind tunnel test train accelerating system which is characterized by comprising supporting frames, U-shaped slide rails, a test bridge span model, a train model, a special train hook device and a special train tripping device, wherein the supporting frames are mounted at the two ends of the test bridge span model; the supporting frames are provided with the U-shaped slide rails; the U-shaped slide rails are provided with the special train hook device and the special train tripping device; the train model is suspended on the special train hook device; and the special train tripping device is arranged at the contact part of the special train hook device and the train model. The U-shaped wind tunnel test train accelerating system provides technical guarantees to a test research on pneumatic performance of a wind-vehicle-bridge system.

The invention discloses an openable-closeable cabin door unmanned aerial vehicle load lifting mechanism. Ascending and descending of a platform and vertical limiting of a pod mounting plate are realized through two ball screws, vertical movement of the platform and horizontal limiting of the pod mounting plate are realized through two carbon tubes. A motor drives the ball screws to run, a synchronous wheel ensures the ball screws to move synchronously, the ball screws drive a middle mounting plate to lift and lower a task load, a servo motor controls opening and closing of a cabin door through a signal processing system, a Hall sensor and a magnetic block control a lifting mechanism to stop moving, and a lock fastener and a lock block ensure stability of the platform after working and being recycled. The task load is connected with the middle mounting plate and can be replaced at any time. By using the mechanism, closing load of the cabin door is unaffected in the process of launching and recycling an unmanned aerial vehicle, and pneumatic performance of the unmanned aerial vehicle can be ensured. The load can be replaced and mounted at any time before the unmanned aerial vehicle takes off, the mechanism is convenient, quick, small in occupied space, light in weight and economically feasible, and overall performance, safety and reliability of the unmanned aerial vehicle are improved greatly.

The invention discloses an anti-congelation and anti-icing electric control device of a wind driven generator blade and belongs to the technical field of new energy power generation equipment. The anti-congelation and anti-icing electric control device comprises a power supply module, a temperature sensor, a humidity sensor, a control module and a heating module, wherein the control module is used for receiving monitoring signals of the temperature sensor and the humidity sensor and analyzing the monitoring signals; the heating module is used for receiving a control instruction transmitted by the control module and generating heat energy according to the control instruction, wherein signal output terminals of the temperature sensor and the humidity sensor are electrically connected with a digital quantity input terminal of the control module through analog-to-digital converters; and a signal output terminal of the control module is electrically connected with the control module. The anti-congelation and anti-icing electric control device has the advantages as follows: according to the temperature and humidity information on the surrounding of the wind driven generator blade, the heating module is intelligently controlled by using the control module; therefore, congelation and icing phenomenon of the wind driven generator blade can be prevented, the pneumatic property of the wind driven generator blade is guaranteed and the working efficiency of the wind driven generator blade is increased.

The invention relates to a wing for an axial-flow fan with tail-edge jetting in the technical field of impeller, which comprises an inputting pipe for the controlling gas, tail-edge jetting holes, a main gas passageway, branch gas passageways, diversion components, and the wing body. The inner of the wing body is a cavity, the main gas passageway and the branch gas passageways are in the cavity; the inputting pipe for the controlling gas is located outside the wing body; one end of the inputting pipe for the controlling gas is communicated with the main gas passageway, another end is communicated with the branch gas passageways; between the branch gas passageways, diversion components are provided. The tail-edge jetting holes are located at the tail edge of the wing body in equi-spaced way, the diameters of the tail-edge jetting holes are arranged in arithmetic progression along the tail edge of the wing body from the end with an inputting pipe for the controlling gas of the wing body to another end of the wing body. The invention fills the speed loss of the tail edge of the wing body, reduces the speed pulse at the tail area, reduces the interference from the downstream blades, and drops the interference noise.

The invention relates to a three-dimensional inward turning air inlet channel design method capable of controlling shock wave and pressure distribution at the same time. The method comprises the following steps:(1) incoming flow conditions and incident shock wave curves are designed, and wall surface molded lines are reversely solved according to a characteristic line method; (2) extracting pressure distribution of the wall surface molded line obtained in the step (1), and designing downstream pressure distribution; (3) according to the pressure distribution obtained in the step (2), reverselysolving a wall surface molded line by utilizing a characteristic line method; (4) taking the three-dimensional revolution of the wall surface molded line obtained in the step (3) as a bus of a basicflow field, and solving the whole flow field; (5) designing a three-dimensional inward turning air inlet passage inlet in polar coordinates, and performing streamline tracking in the basic flow fieldobtained in the step (4) to obtain a three-dimensional inward turning air inlet passage compression profile; and (6) the three-dimensional inward-turning air inlet channel compression molded surface obtained in the step (5) is stretched in the equal-straight-direction backward direction to obtain an isolation section. The advantages of the three-dimensional inward turning air inlet channel are kept, meanwhile, the incident shock wave and wave back isentropic compression process is designed, and therefore a new degree of freedom is introduced for design of the three-dimensional inward turning air inlet channel.

The invention discloses an aerodynamic layout and a design method of an air-breathing hypersonic aircraft with a high lift-drag ratio, and belongs to the field of aerodynamic layout design of aircrafts. A front body pre-compression face, a first-stage outer compression face, a second-stage outer compression face, a third-stage outer compression face, a stamping runner, a first exhaust nozzle and a second exhaust nozzle are arranged in the machine body/inner runner pneumatic layout from front to back, an air inlet channel splitter plate is located in the second-stage outer compression face, and a turbine runner is connected to the upper portion of the stamping runner in parallel. The aircraft has efficient high-speed cruise flight capability, meets the high lift requirement of the aircraft in a low-speed section, solves the problem of flow separation of high-speed wings in a low-speed take-off state, and can ensure the wide-speed-range flight performance of the aircraft.

The invention discloses a wind turbine blade capable of improving stalling characteristics based on fractal optimization. The wind turbine blade comprises a blade body positioned at the front part corresponding to a flow incoming direction, and a blade tail edge positioned at the back part corresponding to the flow incoming direction; the blade tail edge is distributed with multiple elastic pieces and fractal holes with fractal structures; the elastic pieces are arranged on suction surface sizes of the fractal holes, and are suffered from a pressure difference effect of a blade pressure surface and a suction furnace to locate in a closing or opening state; and the fractal holes penetrate through the pressure surface and the suction furnace. When the blade incoming flow attack angle dose not reach stalling conditions, as the pressure difference between the blade pressure surface and the suction surface is smaller, and at this time, the elastic pieces are located in the closing state to stop passing of a fluid to stop small-scale turbulence so as to guarantee the blade pneumatic performances; when the blade incoming flow attack angle is increased to reach the stalling conditions, the elastic pieces are located in the opening state due to the effect of pressure difference, and the fluid can pass through the fractal holes to generate small-scale turbulence on the suction furnace to delay separation of a blade surface boundary layer, so that the formation of stalling clusters is delayed, and the purpose of improving the blade stalling characteristics is achieved.

The invention discloses a novel noise reduction worm tongue structure for centrifugation. The worm tongue structure comprises a worm tongue main body, and a cavity structure positioned under the wormtongue main body; the worm tongue main body comprises a worm tongue bent part, and a cavity wall positioned under the bent part; the inner side surface of the worm tongue bent part is connected with the end surface of the cavity wall to form a cavity structure; the left and right ends of the worm tongue bent part are connected with outlet side walls of a worm shell; the left and right ends of thecavity wall and the outlet side walls of the worm shell form a closed cavity; and a through hole is formed in the bent place of the worm tongue bent part. A noise reduction worm tongue through hole ismatched with the cavity structure to precisely control the flowing of air near the worm tongue to balance the flow distribution of air currents near the worm tongue in an outlet of a fan and the pressure distribution of the air currents near a low-pressure side, so that the pressure of the air currents near the worm tongue is more uniform, the pressure pulsation is reduced, and the rotating noiseis lowered; and meanwhile, the air currents near the worm tongue are adjusted through the through hole, so that eddy currents near the worm tongue can be reduced, and the eddy noise of the fan is lowered.

The invention discloses an inner-bulge S-bend air inlet channel with a middle control section and a method, and belongs to the technical field of aircrafts. The inner-bulge S-bend air inlet channel comprises an air inlet channel first S-bend section and an air inlet channel second S-bend section, and the interface of the air inlet channel first S-bend section and the air inlet channel second S-bend section is the middle control section. The middle control section serves as an outlet of the first S-bend section and also serves as an inlet of the second S-bend section. The middle control section is inwards concave and is composed of an upper arc, a lower arc, a left arc and a right arc which are tangent, the upper arc and the lower arc have the same circle center and central angle, and the circle center is located above the middle control section. The upper section arc of the middle control section is concave downwards towards the center line of the air inlet channel, so that an inner bulge structure is formed in the position, corresponding to the middle control section, of the upper inner surface of the air inlet channel. The inner bulge structure is introduced, air flow reversely deflects twice in a pipeline through the middle control section, all-directional shielding of components in the air inlet channel and compressor blades is achieved, and the air inlet channel has the good radar stealth performance when the center offset distance of the inlet face and the outlet face is small.

The invention belongs to the field of structural designs of aircraft engines, and in particular relates to an asymmetrical exhaust structure of a lift fan. The asymmetrical exhaust structure comprisesan asymmetrical caudal vertebra (4), an adapting cartridge receiver (5) and a sprayer nozzle (6). A fairing adapting section is arranged on the adapting cartridge receiver (5), wherein one end of thefairing adapting section is connected to the asymmetrical caudal vertebra (4) and is coaxial with the lift fan; and the other end of the fairing adapting section is connected to the sprayer nozzle (6) and is coaxial with the sprayer nozzle (6). The middle part of the fairing adapting section is a smooth curved surface section. The lift fan is of an asymmetrical exhaust structure. The adapting cartridge receiver provided with the fairing adapting section is connected to the asymmetrical caudal vertebra and the sprayer nozzle, so that the exhaust loss of the lift fan can be reduced and the sprayer nozzle performance and the fan pneumatic performance are guaranteed.

The invention discloses an anti-icing structure of an inlet guide vane of an aero-engine and the inlet guide vane. The anti-icing structure comprises a mounting part and a deformation part. The inlet guide vane comprises a guide vane body, and the mounting part is arranged on the surface of the guide vane body. The deformation part is arranged in the mounting part. When the temperature around the aero-engine is lower than a first temperature, the deformation part is increased from a first volume to a second volume. When the temperature around the aero-engine is higher than a second temperature, the deformation part is reduced from the second volume to the first volume. By the adoption of anti-icing structure, ice on the surface of the guide vane body can be broken due to deformation of the surface of the guide vane body to which the ice is attached, and the deicing effect is achieved; the influence on the structural strength of the inlet guide vane can be reduced, and the aerodynamic performance of the inlet guide vane is kept; and the volume change of the deformation part is a reversible and repeatable process, so that the part, easy to freeze, of the inlet guide vane can be deiced according to the temperature change.

The invention discloses an aviation blade profile chord length detection method and system based on secant rotation iteration, and belongs to the field of aviation blade profile detection, and the method comprises the steps: taking a connection line between two farthest points in a blade profile contour point set as a secant, rotating or translating the secant to enable the secant to intersect with a blade basin, and obtaining a leading edge point set and a trailing edge point set; rotating the secant in the range of the leading edge point set and the trailing edge point set to obtain tangentpoints of the secant and the leading edge and the trailing edge, and calculating the distance between the two tangent points to obtain the chord length. The chord length calculated through the methodis high in accuracy, the size of the blade can be controlled through the chord length of the blade, meanwhile, if the chord length is too short, the strength of the blade is reduced, the impact resistance and the aerodynamic performance of the blade are affected, and therefore the aerodynamic performance and safety of the blade can be guaranteed when the chord length of the blade is detected through the method.

The invention provides a chord length adaptive stretching and contracting type flapping wing suitable for an active torsion flapping mechanism and a flapping wing machine. The chord length adaptive stretching and contracting type flapping wing comprises a front beam, a rear beam, stretching and contracting type wing ribs and a lap-joint type outer panel skin, wherein a flapping wing skeleton is distributed and provided with the multiple stretching and contracting type wing ribs in the direction from a wing root to a wing tip, each stretching and contracting type wing rib comprises a front edgesegment wing rib, a rear edge segment wing rib and a stretching and contracting sleeve, and the lap-joint type outer panel skin comprises a front edge segment outer panel skin and a rear edge segmentouter panel skin. The chord length adaptive stretching and contracting type flapping wing suitable for the active torsion flapping mechanism and the flapping wing machine have the advantages that (1), the chord length adaptive stretching and contracting type flapping wing has the ability of changing the wing area adaptively to be capable of adapting to a certain degree of chordal stretching and contracting generated by the active torsion flapping mechanism, the aerodynamic shape integrity of an active torsion flapping wing in the whole flapping process is guaranteed, thus the aerodynamic performance is guaranteed, and the flapping wing flight aerodynamic efficiency is improved; and (2), by means of design on the outer panel skin, the wing face is guaranteed to be a whole with a complete aerodynamic shape in the flapping process, and then the flapping wing overall aerodynamic characteristic is guaranteed.

The invention relates to a telescopic folding device and a wing, the device comprises a telescopic mechanism and a folding mechanism, the telescopic mechanism comprises a movable end rib, at least four guide rail assemblies, at least two first drivers and at least two second drivers, the first guide rail assembly and the second guide rail assembly are arranged at one end of the movable end rib, the third and fourth guide rail assemblies are arranged at the other end of the movable end rib; the first drivers are connected with the first and second guide rail assemblies; the second drivers are connected with the third and fourth guide rail assemblies, and the movable end rib is driven by the at least two first drivers and the second drivers to do reciprocating motion; and the folding mechanism is connected with a movable end rib of the telescopic mechanism. The folding wing structure is designed, the wing tip structure folding function is achieved, meanwhile, it is guaranteed that gaps do not appear on the upper face and the lower face of the wing before folding, it is avoided that a complex sealing structure is designed due to large gaps, the completeness of the upper wing face and the lower wing face is guaranteed, and the aerodynamic performance of an aircraft is guaranteed.

The invention discloses a flame tube with a novel cooling structure. The flame tube comprises a combustion chamber casing, wherein a flame tube body and a large elbow are arranged in the combustion chamber casing, an inlet diffuser is arranged on one side of the combustion chamber casing, a lap joint section is arranged in the combustion chamber casing and between the flame tube body and the large elbow, an outer ring main combustion mixing hole is formed in an outer ring of the flame tube body, a head swirler mounting hole is formed in a head ring of the flame tube body, an inner ring main combustion mixing hole is formed in an inner ring of the flame tube body, the flame tube body comprises the outer ring, the head ring, the inner ring and a small elbow, the outer ring, the head ring, the inner ring and the small bent pipe of the flame tube body are of an integrated structure, and zigzag small-aperture circular cooling channels are formed in the wall faces of the outer ring, the head ring, the inner ring and the small elbow of the flame tube body and the wall face of the large elbow respectively, so that the number of holes formed in the wall faces is greatly reduced, cracks of the holes are reduced, and the service life of the flame tube body is prolonged.

The invention discloses a composite protection film of a wind power blade and a detachable wind power blade protection device. The composite protection film of the wind power blade comprises a supporting layer and an elastic layer, wherein the supporting layer is made of a material with certain strength, and the elastic layer comprises one of a thermoplastic elastic layer, a rubber layer and a foaming layer. The composite film, first fixing ribs, second fixing ribs, grid plates and a skin jointly form the detachable wind power blade protection device. Raw material particles of the elastic layer and raw material particles of the supporting layer are extruded together through a special mold head, so that the composite protection film is obtained. The bonding force between the two layers of the composite film is high, the two layers cooperate with each other to play a role, and the wind power blade is effectively protected.