180results about "Influencers by generating vortices" patented technology

The invention provides a heterogeneous micro-costal membrane and a preparation method thereof.The heterogeneous micro-costal membrane comprises a front face and a back face, the front face is provided with a vortex generator array and a rib array in a protruding mode, the vortex generator array and the rib array are arranged in a continuous mode, the vortex generator array comprises a plurality of vortex generators which are arranged in a tiled mode, and the rib array is arranged on the back face. The rib array comprises a plurality of ribs which are arranged in a tiled mode, the rib array sequentially comprises a first protection layer, a second protection layer, a supporting layer, a back glue layer and a bottom film layer from the front face to the back face, the first protection layer is made of a super-hydrophobic modified material, the second protection layer is formed by bombarding a chromium target through argon ions and depositing chromium atoms, and the supporting layer is made of a super-hydrophobic modified material. The supporting layer is made of an epoxy resin material, and the back glue layer is prepared by performing corona treatment on the surface of the supporting layer and coating glue; the bottom film layer can be removed. The heterogeneous micro-costal membrane provided by the invention has relatively high overall strength, wear resistance, corrosion resistance and a remarkable drag reduction effect.

An aircraft including an aircraft wing wherein the aircraft wing includes a fixed wing, a wing tip device at an end of the fixed wing, and a wing tip device actuator configured to move the wing tip device between a flight configuration and a ground configuration. The wing tip device actuator includes a shaft which is rotatable about a shaft axis to selectively displace respective first and second elbows of the wing tip device actuator towards or away from the shaft. The first elbow is connected to the wing tip device and the second elbow is connected to the fixed wing such that movement of the respective elbows away from the shaft causes the wing tip device to move away from the flight configuration and such that movement of the respective elbows towards the shaft causes the wing tip device to move towards the flight configuration.

The present invention provides an extreme STOL airplane comprising: a fuselage and a wing having port and starboard wing sections. Each wing section has a body with top and bottom surfaces, leading and trailing edges, and a span, i.e., the distance between a wing section tip and root connected to the fuselage, and flap coves. Aerodynamic elements integrated into the wing sections include Fowler flaps, having leading and trailing edges and a span, flap tracks, wherein the flap tracks are external to the wing section body, extend aftward beyond the trailing edge of the wing section, and are configured to enable the Fowler flaps to rotate or deflect and to translate or extend and retract, Frise ailerons, wherein the Frise ailerons are located outboard of the Fowler flaps, and spoilerons located over the leading edge of the Fowler flaps when the flaps are in the fully extended position.

An aircraft wing is disclosed having a fixed wing with a tip, and a wing tip device rotatably mounted on a hinge at the tip of the fixed wing, such that the wing tip device is rotatable about the hinge, and an actuation system for rotating the wing tip device about the hinge. The actuation system includes a motor, at least one geared rotary actuator, a reduction gearbox, a clutch for selectively decoupling rotation of the motor from rotation of the geared rotary actuator, the geared rotary actuator is driveable by the motor and arranged to convert rotary motion into a different rotary motion and is arranged to rotate the wing tip device relative to the tip of the fixed wing. The clutch is coupled to the geared rotary actuator by a shaft and the shaft is fixed in rotation to a mass disposed radially outwardly from the shaft.

A wing assembly for an aircraft including a fixed wing and a wing tip device rotatable about a joint assembly. The wing assembly is configurable between a flight configuration for use during flight and a ground configuration for use during ground-based operations. The wing assembly includes joint-bordering structure located adjacent the joint assembly, the joint-bordering structure having a rib, an upper cover and a lower cover. The joint assembly includes a moveable panel. The panel includes a first hard stop element, and the joint-bordering structure includes a second hard stop element. The first hard stop element is integrally formed in the panel, and the second hard stop element is integrally formed in one of the rib, the upper cover or the lower cover of the joint-bordering structure. The hard stop elements may be configured to self-align.

A method for reducing aerodynamic loads on an aircraft located on the ground includes moving a wing tip relative to a fixed wing to an on-ground configuration in which the span of the wing is reduced. The wing tip includes an airflow channel, which is in an open configuration in which airflow through the channel between upper and lower surfaces of the wing tip is enabled, when the wing tip is in the on-ground configuration. This has been found to reduce aerodynamic loads on the wing tip and / or parts of the aircraft connected thereto during this phase of the aircraft's use.

An aircraft wing assembly (31) is disclosed. The aircraft wing assembly (31) comprises a wing tip device (11) rotatable about a hinge axis (H) between a flight configuration and a ground configuration and a wing tip device actuation system (20) comprising a bell crank (21) and a linear actuator (23). The bell crank (21) is rotatably mounted to the aircraft wing assembly (31) and is rotatable about a bell crank axis (C) which is offset from the hinge axis (H). The linear actuator (23) is operable to rotate the bell crank (21) about the bell crank axis (C) to thereby move the wing tip device (11) about the hinge axis (H), between the flight configuration and the ground configuration.

A wing (3) for an aircraft (1) including a fixed wing (5), a foldable wing tip portion (9) mounted to the fixed wing (5) via a hinge (11) in a manner rotatable about a hinge axis (13) between an extended position and a folded position, an actuation device (24) for actuating the foldable wing tip portion (9) for movement between the extended position and the folded position, and an arresting device (37) for arresting the foldable wing tip portion (9) in the extended position and / or in the folded position. The arresting device (37) includes a clutch device (39) configured for selectively holding the foldable wing tip portion (9) in a fixed position relative to the fixed wing (7) and releasing the foldable wing tip portion (9) for movement relative to the fixed wing (7) about the hinge axis (13).

Various control surfaces are disclosed that can cause a side force along a lateral axis running through a center of mass of a craft when the control surfaces are deployed in a same direction. The side force can be created without creating a yawing moment on the craft. The control surfaces can be on a winglet of the craft. Also disclosed are a split-flap control surface and stand-alone ailerons.

A hinge arrangement for an aircraft wing is provided, for providing a folding wingtip. The hinge arrangement comprises a first mount for securing in a first wing section and a second mount for securing to a second relatively foldable wing section, each having a plurality of hinge elements arranged along a hinge axis. The arrangement comprises a latching arrangement for securing the mounts against folding movement, the latch arrangement having a latch axis being spaced from the hinge axis in a substantially spanwise direction for example so as to extend into the wing box. This can provide an efficient arrangement which may be contained within the outer mould line of the wing which can reduce the need for a fairing.

The present invention provides an extreme STOL airplane comprising: a fuselage and a wing having port and starboard wing sections. Each wing section has a body with top and bottom surfaces, leading and trailing edges, and a span, i.e., the distance between a wing section tip and root connected to the fuselage, and flap coves. Aerodynamic elements integrated into the wing sections include Fowler flaps, having leading and trailing edges and a span, flap tracks, wherein the flap tracks are external to the wing section body, extend aftward beyond the trailing edge of the wing section, and are configured to enable the Fowler flaps to rotate or deflect and to translate or extend and retract, Frise ailerons, wherein the Frise ailerons are located outboard of the Fowler flaps, and spoilerons located over the leading edge of the Fowler flaps when the flaps are in the fully extended position.

The application discloses a vortex generator with adjustable interval and angle, and relates to the technical field of vortex generators.The vortex generator with adjustable interval and angle comprises a plurality of vortex generator units, an interval adjusting assembly and an angle adjusting assembly;each vortex generator unit comprises a flow guide plate, a support rod, an interval adjusting sliding block and a rotary rocker arm;one end of each support rod is connected and fixed with the corresponding flow guide plate, and the other end penetrates through a through hole and is rotationally connected with the corresponding interval adjusting sliding block;one end of the rotary rocker arm is connected and fixed with the corresponding support rod;the interval adjusting assembly can drive the plurality of interval adjusting sliding blocks to slide so as to adjust the interval between two adjacent flow guide plates;and the angle adjusting assembly can drive the plurality of rotary rocker arms to rotate around the axial direction of the support rod so as to adjust the angle of the flow guide plate.The application can automatically adjust the interval and angle of the vortex generator, so that the vortex generator can be adapted to more complex air flow environments and the energy efficiency ratio can be optimized.

A winglet system is configured to be secured to a wing of an aircraft. The winglet system includes an upper blade, a lower blade, and a cover plate integrally formed with the upper blade and the lower blade. The cover plate covers an opening within a transition joint between the upper blade and the lower blade. The cover plate remains coupled to the upper blade if the lower blade is separated from the winglet system.

Vortex generators include a vane that rotates between a stowed position and a deployed position. In the deployed position, the vane is configured to reduce noise created by an airflow around or along the aerodynamic surface. A spring controls deployment of the vane and includes an elongate tube formed of a superelastic shape memory alloy. To deploy or stow the vane, the elongate tube is twisted due to a stress-induced response caused by airflow around or along the aerodynamic surface, thereby building up and storing torque in the elongate tube. A mounting base receives the vane when it is in the stowed position, and operatively couples the vortex generator to the aerodynamic surface. To create the torque in the elongate tube, a first end of the elongate tube is fixed with respect to the aerodynamic surface while a second end of the elongate tube is free to rotate.

An aircraft wing (5, fig 1) and tuning the wing comprises a fixed wing (7) and a wing tip device (9) rotatably mounted on a hinge 11 at a tip of the fixed wing. An actuation system 20 rotates the wing

The invention discloses a vortex generator with adjustable spacing and angle, and relates to the technical field of vortex generators. The vortex generator with the adjustable distance and angle comprises a plurality of vortex generator units, a distance adjusting assembly and an angle adjusting assembly. Each vortex generator unit comprises a flow guide plate, a supporting rod, a distance adjusting sliding block and a rotating rocker arm. One end of each supporting rod is fixedly connected with the corresponding flow guide plate, and the other end of each supporting rod penetrates through the corresponding via hole and is rotationally connected with the corresponding interval adjusting sliding block. One end of each rotary rocker arm is fixedly connected with the corresponding supporting rod; the distance adjusting assembly can drive the multiple distance adjusting sliding blocks to slide so as to adjust the distance between every two adjacent flow guide plates. The angle adjusting assembly can drive the multiple rotating rocker arms to rotate in the axial direction of the supporting rod so as to adjust the angle of the flow guide plate. The space and the angle of the vortex generator can be automatically adjusted, so that the vortex generator is adaptive to more complex airflow environments, and the energy efficiency ratio is optimized.

The present invention provides a heterogeneous micro-rib membrane and a preparation method thereof, comprising a front and a back surface, wherein the front surface is provided with a vortex generator array and a rib array, wherein the vortex generator array is arranged in series with the rib array, wherein the vortex generator array includes a plurality of flatly arranged vortex generators, and the rib array includes a plurality of flatly arranged ribs. From the front surface to the back surface, the membrane comprises a first protective layer, a second protective layer, a support layer, a backing layer, and a bottom film layer, wherein the first protective layer is made of a super-hydrophobic modified material, the second protective layer is formed by argon ion bombardment of a chromium target and deposition of chromium atoms, the support layer is made of an epoxy resin material, and the backing layer is prepared by corona treating the surface of the support layer and applying glue; the bottom film layer is removable. The heterogeneous micro-rib membrane provided by the present invention has high overall strength, wear resistance, corrosion resistance, and significant drag reduction effect.

The invention relates to a novel wingtip winglet of an aircraft. The installation angle of the winglet is a forward swept wing, and the winglet is a dihedral winglet. Compared with other winglets, the novel winglet has more excellent aerodynamic characteristics, the vortex phenomenon of the wingtip of a large wing of an aircraft can be reduced, the resistance of the aircraft is reduced, and the lift force of the aircraft is increased.

A wing tip device for fixing to the outboard end of a wing, the wing defining a wing plane, the wing tip device comprising: an upper wing-like element projecting upwardly with respect to the wing plane and having a trailing edge; and a lower wing-like element fixed with respect to the upper wing-like element and having a root chord and a trailing edge, the lower wing-like element root chord intersecting with the upper wing-like element, and the lower wing-like element projecting downwardly from the intersection, wherein the upper wing-like element is larger than the lower wing-like element and the trailing edge of the lower wing-like element is adjacent the trailing edge of the upper wing-like element at the intersection, and wherein an included angle between the upper and lower wing-like elements at the intersection is less than, or equal to, 160 degrees.

An aircraft wing with a moveable leading edge device mounted towards the leading edge of the wing. The leading edge device is moveable between a first configuration, a second configuration and a third configuration. In the first configuration, the leading edge device is retracted fully within the wing profile. In the second configuration, a portion of a surface of the leading edge device is extended away from the wing profile and into the oncoming airflow on the lower surface of the wing as the wing moves through the airflow to increase lift produced by the wing. In the third configuration, a portion of the surface of the leading edge device is extended away from the wing profile into the oncoming airflow over the upper surface of the wing to impair lift produced by the wing.

The invention relates to a turbine engine motor having an unducted fan, the motor comprising a plurality of guide vanes (10) made of composite materials. Each of the guide vanes is intended to be fixed only to an inner fixed structure (20) of the turbine engine. The guide vane (10) comprises: an aerodynamic airfoil (12) which extends radially outwards along a radial axis (X-X); an outer radial end (14); an inner radial end (16); and an inner platform (18) which is configured to be fixed to the inner fixed structure (20), is attached to the inner radial end (16) by an inner attachment elbow (16a) and extends from the inner radial end (16) in a single first tangential direction secant to the radial axis (X-X) of the aerodynamic airfoil (12). The inner platform (18) comprises an opening (18a) for receiving a radial fixing system (30) for independently fixing each of the guide vanes to the inner fixed structure (20). A reinforcement element is arranged radially between the inner platform (18) of each of the guide vanes (10) and the inner fixed structure (20).

The invention discloses a coronal turbulence drag reduction micro-rib structure, belongs to the technical field of aviation engineering, and aims to improve the drag reduction effect of the micro-rib structure. Comprising a surface base body and a plurality of small resistance reduction ribs, the small resistance reduction ribs are of strip structures with uniform sections, extend in the flow direction of fluid and are sequentially arranged at intervals in the direction perpendicular to the flow direction of the fluid, each small resistance reduction rib comprises a crown body and a rib body which are vertically connected, and the cross section of each crown body and the cross section of each rib body are each in an isosceles trapezoid shape. The lower bottom face of the crown body is aligned with the upper top face of the rib body, an isosceles trapezoid groove is formed in the upper top face of the crown body and penetrates through the crown body in the length direction of the crown body, two protruding structures which are symmetrical left and right are formed in the top of the crown body, and the lower bottom face of the rib body is connected with a surface base body. The pipe body can generate a stronger secondary vortex structure under the same flow field condition, and has a stronger inhibition effect on spanwise oscillation of a main flow to a vortex, so that the resistance reduction effect is improved.

Vortex generators include a vane that rotates between a stowed position and a deployed position. In the deployed position, the vane is configured to reduce noise created by an airflow around or along the aerodynamic surface. A spring controls deployment of the vane and includes an elongate tube formed of a superelastic shape memory alloy. To deploy or stow the vane, the elongate tube is twisted due to a stress-induced response caused by airflow around or along the aerodynamic surface, thereby building up and storing torque in the elongate tube. A mounting base receives the vane when it is in the stowed position, and operatively couples the vortex generator to the aerodynamic surface. To create the torque in the elongate tube, a first end of the elongate tube is fixed with respect to the aerodynamic surface while a second end of the elongate tube is free to rotate.

To transition airflow from laminar to turbulent locally within an engine nacelle inlet.SOLUTION: A nacelle inlet 21 includes an inner flow surface, an outer flow surface, and a turbulator segment 50 positioned along the inner flow surface. The turbulator segment extends outward beyond the inner flow surface 61 and is configured to transition laminar airflow along the inner flow surface upstream from the turbulator segment to turbulent airflow downstream from the turbulator segment.SELECTED DRAWING: Figure 4

To eliminate or minimize damage caused by a typhoon.SOLUTION: When a typhoon is expected, at a low-pressure part where a week typhoon in front of the same can be generated, a counterclockwise updraft 3 is induced from near a sea surface 4 by means of a jet injection 2 severely released rearward of a jet plane 1 descending clockwise beforehand from a high altitude to reduce the energy of a typhoon approaching after generating the typhoon; therefore, if needed, repeatedly generating a week typhoon enables a sufficient reduction of the energy of the typhoon itself.SELECTED DRAWING: Figure 1

Reduction of the induced drag by placing a convergent-divergent nozzle at the wingtip which creates an additional airflow with a velocity greater than the airflow velocity around the wing, which changes the nature and direction of the airflow along the upper and lower surfaces at the wingtip, thus forming a new vortex of lesser intensity, which is shifted to the sides of the wingtip, creating conditions for better airflow, less induced drag, and greater aerodynamic force.

The present invention provides a method for reducing heat and drag on the wall of an aircraft based on an aerodynamic metasurface. An aerodynamic metamaterial metasurface is provided on the surface of the aircraft for reducing heat and drag. The metamaterial substrate of the aerodynamic metamaterial metasurface adopts a porous medium material and is densely covered with micro-nano pores that can allow cooling gas inside the aircraft to seep out. Microstructured surface units are periodically arrayed on the surface of the metamaterial substrate, and the microstructured surface units have concave cavities. When an ultra-high-speed incoming flow passes through the aerodynamic metamaterial metasurface, part of the incoming flow gas stays in the concave cavities to form trapped vortices, thereby reducing the shear effect between the incoming flow and the surface of the metamaterial substrate, thereby playing a drag reduction role. The cooling gas seeps out from the surface of the metamaterial substrate from the inner side of the bottom surface of the metamaterial substrate through the micro-nano pores inside the metamaterial substrate, forming a cooling gas film on the surface of the metamaterial substrate, reducing the friction between the mainstream and the surface of the metamaterial substrate, and isolating the heat exchange between the high-enthalpy mainstream and the surface of the metamaterial substrate, thereby playing a heat reduction role.

The invention discloses a power device of a rear-mounted aileron variable-pitch ducted propeller aircraft. The power device comprises a power part, a variable-pitch execution part, a duct, ailerons and a variable-pitch driving part, a power part coaxial with the duct is arranged in the duct; the power part is coaxially provided with a variable-pitch driving part, the driving end of the variable-pitch driving part is provided with a variable-pitch execution part, and the variable-pitch execution part is connected with the power part; the variable-pitch driving part is used for driving a paddle mounted on the variable-pitch execution part to adjust a variable-pitch angle, and the power part is used for driving the paddle to axially rotate around the power part; one end, far away from the power part, of the duct is provided with an extension part mounting aileron. Variable pitch angle adjustment of the blades can be achieved, power can be provided efficiently in different flight processes and at different flight speeds, noise can be reduced through the ducts, upward control force in the longitudinal direction of the wings is provided by enabling wake flow of the propeller blades to flow through the surfaces of the ailerons, and the flight speed is increased. And the aircraft power device keeps high flight efficiency in the whole flight process and has the capability of diversification of manipulation control.

Reducing aeroelastic flutter in a wing tip 404 comprises providing at an end of a fixed wing 402 a rotatable wing tip rotatable a first distance to lower the centre of gravity of the wing tip, i.e. a