352results about "Heat reducing structures" patented technology

The invention discloses a wing structure with flow control and separation control of laminar flow. A plurality of A micropores are arranged at the front end of the vertical center line of the upper wing surface of a wing, and are neatly arranged in lines and rows to form an inspiration zone; a plurality of B micropores are arranged at the back end of the vertical center line of the upper wing surface of the wing, and are neatly arranged in lines and rows to form an air blowing zone; a communicating channel in the inspiration zone and the air blowing zone is called an air flow channel, namely, an inlet of the air flow channel is communicated with the inspiration zone, while an outlet of the air flow channel is communicated with the air blowing zone; the air flow channel is arranged in the wing; the air flow channel consists of a front air flow channel, a middle air flow channel and a back air flow channel, wherein, the middle air flow channel is in an elliptic shape; an inspiration pump is arranged in the middle air flow channel; the front air flow channel and the back air flow channel are designed into a flaring shape and are oppositely arranged. The wing of the invention can delay the transition of a bounding layer from lamina flow to turbulent flow by a control method for adjusting the capacity of blowing and inspiration so as to reduce the frictional resistance on an object surface, and can delay the flow separation so as to improve the stall performance of wing profile.

The invention relates to a self-adapting morphing trailing edge based on SMA, belonging to a self-adapting morphing wing which combines intellectual materials and structures. The wing is divided into 2-5 trailing edge segments, the adjacent trailing edge segments are connected by joints (5, 6 and 7) arranged at wing ribs, a deflection driving mechanism is arranged between the adjacent trailing edge segments, and the deflection of the entire trailing edge is realized by accumulative effects of the trailing edge segments. The self-adapting morphing wing is characterized in that the deflection driving mechanism comprises an upper SMA wire (10), a lower SMA wire (11) and a current excitation unit (15) which are respectively connected with the adjacent trailing edge segments. The self-adapting morphing trailing edge has the advantages of simple structure and easy control and achieves the goal of quick, stable and accurate pre-deformation of the trailing edge.

The invention provides a low-Reynolds-number airfoil profile with cooperative fluidic control, and a control method thereof. The low-Reynolds-number airfoil profile with cooperative fluidic control is characterized in that an air jet port (2) is formed in the front edge of the upper surface of an airfoil profile body (1); an air suction port (3) is formed in the rear edge of the upper surface of the airfoil profile body (1); the air jet port (2) is communicated with the air suction port (3) through an air flow pipeline (5) which is arranged inside the airfoil profile body (1); an air pump (4) which is used for driving air suction and air jet to be carried out simultaneously is mounted in the air flow pipeline (5); moreover, the air jet port (2) and the air suction port (3) are perpendicular to the upper surface of the airfoil profile body (1) respectively. The lift of the airfoil profile is greatly increased, the resistance is obviously reduced, and the stalling characteristic of the airfoil profile is improved, so that the aim of efficiently improving the aerodynamic performance of an aircraft is achieved; in addition, the low-Reynolds-number airfoil profile also has the characteristic of low energy consumption.

The invention relates to a wingtip eddy diffusion device, comprising an upper winglet and a lower winglet which play the role of an end plate to block bottom wing surface air current from flowing to the upper wing surface; the upper winglet and the lower winglet are both of symmetrical airfoil profile; the symmetrical airfoil profile can avoid intense shock wave under supercritical air current condition, extra wave drag; certain deflection exists in wingtips of corresponding wings of the two winglets, thus improving lifting force of wings. The invention has the advantages of weakening wingtip eddy diffusion and trailing vortex intensity, downwash field of wingtips of wings and induced drag.

A hinge sealing element to substantially prevent airflow between an edge of a moveable aero surface structure and an edge of a fixed skin of an aircraft wing in the vicinity of the ribs to which the moveable aero surface structure is pivotally mounted is disclosed. The hinge sealing element comprising a deformable elongate, resiliently flexible strip having an upper face to contact a trailing edge of a fixed skin and one end attachable to a moveable aero surface structure such that said strip extends from an upper surface of said moveable aero surface structure beyond its leading edge and into a space between the edge of the fixed skin and the rib of the aircraft wing to a second, free end. The strip is configured such that it deforms during pivotal movement of the moveable aero surface structure to maintain contact between the upper face of the flexible element and the trailing edge of the fixed skin.

The invention provides a sharkskin-imitating resistance-reducing wing and belongs to the field of resistance-reducing wings. By the adoption of the sharkskin-imitating resistance-reducing wing, the problem that flying resistance is formed by vortexes on the surface of a plane is solved. According to the sharkskin-imitating resistance-reducing wing, a speed sensor (3) is installed on the surface of a plane body (1); a skin layer (4) adheres to the surface of a wing body (2); a set of arc-shaped protruding ribs (5) are arranged on the surface of the skin layer (4); and a groove (6) is formed between every two adjacent protruding ribs (5). A resistance reducing method of the sharkskin-imitating resistance-reducing wing comprises the steps that if the resistance measured by the speed sensor (3) is excessively large, the skin layer (4) is heated so that the temperature of the skin layer (4) can be increased, and a sharkskin-imitating rib structure is formed by means of the tensile force of a variable-trailing-edge wing body or a variable-chord wing body. According to the sharkskin-imitating resistance-reducing wing, the flow field characteristics around the plane are changed through the sharkskin-imitating rib structure, so that the resistance borne by the plane in the flying process is reduced.

The invention provides a low-Reynolds-number airfoil section with multi-seam synergetic jet flow control and a control method. The low-Reynolds-number airfoil section with multi-seam synergetic jet flow control comprises a jet exhaust (2) arranged on the front edge of the upper surface of the airfoil section (1), wherein a suction area (3) formed by a plurality of orderly arranged suction tiny holes (10) is arranged on the rear edge of the upper surface of the airfoil section (1); the jet exhaust (2) is communicated with the suction area (3) through an airflow pipeline (5) in the airfoil section (1) to form a suction loop; an air pump (4) for driving the synchronous operation of suction and injection is installed in the airflow pipeline (5); the jet exhaust (2) and the suction tiny holes (10) are vertical to the upper surface of the airfoil section (1). The suction control technology is applied to the low-Reynolds-number airfoil section, the lifting-resistance characteristic of the airfoil section is improved and the aerodynamic characteristic of a high-altitude vehicle is perfected by controlling the laminar separation of the low-Reynolds-number airfoil section; the airfoil section has the advantages of low power consumption, so that the aerodynamic efficiency of the high-altitude vehicle is improved.

A method for applying an aerodynamically functional coating to an exterior surface is provided. The method includes preparing an uncoated substrate surface to obtain a prepared substrate surface; applying a primer on the prepared substrate surface to obtain a primed substrate surface with a primer layer; applying and curing a basecoat field color to obtain a basecoat field color layer; masking for a basecoat design color on the basecoat field color layer; applying and curing the basecoat design color to obtain a first basecoat design color layer; masking for, applying, and curing any subsequent basecoat design color layers on the basecoat field color layer; applying and curing a clearcoat to obtain a first clearcoat layer; mechanically aero-contouring a plurality of coating edges to obtain an aero-contoured surface; and, applying additional clearcoat to obtain one or more additional clearcoat layers.

The invention discloses a flying device which comprises a device main body, a sealed inner shell located on the device main body and an outer shell located outside the inner shell. An air channel is arranged between the inner shell and the outer shell; the outer shell located on the upper part of the device main body comprises a plurality of spoilers, and a guide inlet and a guide outlet which are communicated with the outside are arranged between every two spoilers or on each spoiler per se. The air channel is communicated with the guide inlet and the guide outlet respectively; the surface path passed by fluid on the external surfaces of the spoilers is longer than the air path in the air channel on internal surfaces of the spoilers in a corresponding zone, therefore, a high pressure zone produced by low flow rate in the air channel is transferred to a low pressure zone produced by high flow rate outside the outer shell from the guide outlet; and the surface path passed by fluid on the outer shell part at the upper part of the wing is longer than the surface path of the outer shell part at the lower part of the wing in the flying direction. The invention can increase the speed ofthe flying device and can save energy.

The invention discloses a front edge impact, micro through passage and air film cooling structure of a hypersonic vehicle. The structure is characterized by comprising an impact cavity, impact holes, micro through passages and air film holes, wherein the impact cavity is positioned in a front edge of a high-speed vehicle; a row of impact holes are formed at axis positions on the middle part at one side of the impact cavity close to a vehicle body; two ends of the impact holes are communicated with the impact cavity and an air supply cavity; the two micro through passages are formed at positions of the impact cavity close to corresponding impact holes on the upper surface and the lower surface of the surface of a wedged body of the vehicle; the impact holes and the corresponding micro through passages are positioned in a plane; and the outer wall of each micro through passage is provided with the air film holes. The structure can be used for cooling a region with high heat flux density, so that the problem that a local high temperature region cannot be effectively cooled due to space limitation is substantially solved.

The invention provides a low-Reynolds-number wing section matched with a full-wing solar unmanned aerial vehicle. The low-Reynolds-number wing section is characterized in that the relative thickness of the wing section is 11%-13%, the maximum camber of the wind section is 2%-4%, and a chordwise position with the maximum camber is 25%-30%; the wing section has a single bent outline in a 70% chordwise range from a front edge. With a full-aircraft matching design, the checking of pneumatic characteristics of a pair of wing sections is finished by checking the corresponding pneumatic characteristics of the wing sections and applying the two wing sections to the full-wing solar unmanned aerial vehicle, and the low-Reynolds-number wing section shows good pneumatic characteristics and engineering enforceability and can meet application demands on low Reynolds number, high lift force and lift-to-drag ratio and large relative thickness of the full-wing solar unmanned aerial vehicle.

A hypersonic aircraft includes one or more leading edge assemblies that are designed to cool the leading edge of certain portions of the hypersonic aircraft that are exposed to high thermal loads, such as extremely high temperatures and / or thermal gradients. Specifically, the leading edge assemblies may include an outer wall tapered to a leading edge or stagnation point. A coolant supply provides a flow of cooling fluid to a porous tip that is joined to the forward end of the outer wall and defines variable porosity and / or internal barriers to direct a flow of cooling fluid to the regions of the leading edge experiencing the highest thermal loading.

The invention discloses an air film and micro straight channel cooling structure for a front edge of a hypersonic vehicle, which comprises an air film cavity, air film holes and micro straight channels. The air film cavity is positioned inside the front edge of the hypersonic vehicle; the air film holes which are arranged in a stagger mode are formed near stagnation points of the front edge at the periphery of the air film cavity; two rows of micro straight channels are formed next to both upper and lower surfaces of a wedge-shaped body of the hypersonic vehicle; and the micro straight channels are communicated with the air film cavity and a high pressure water tank. On the premise of not changing the pneumatic appearance of the air vehicle, the air film cooling mode is combined with the micro channel cooling mode, so that the heat exchange efficiency is effectively improved.

Adaptive structures, systems incorporating such adaptive structures and related methods are disclosed. In one embodiment, an adaptive structure is provided which includes a first structure and at least one microstructure associated with the first structure. The at least one microstructure may include a microscale beam configured to be displaced relative to the first structure upon the adaptive structure being exposed to a specified temperature. The beam may be formed for example, of a metallic material, of multiple different metallic materials, or of a shape memory alloy. In one embodiment, a plurality of the adaptive structures may be associated with micropores of a skin panel. The adaptive structures may be utilized to control the flow rate of a coolant or other fluid through the micropores responsive to a sensed environmental parameter such as, for example, temperature.

The invention relates to an aerodynamic high-performance aerofoil for an aircraft. A leading edge of the aerodynamic high-performance aerofoil is provided with a cavity and an elastic skin covering an opening of the cavity. Through forced vibration loaded on an elastic structure or self-excited vibration produced by strong fluid-structure interaction with fluid in the flight, the aerodynamic high-performance aerofoil can control flow thereby improving aircraft aerodynamic performances. The aerodynamic high-performance aerofoil has the advantages of obvious lift-rising and drag reduction effects, simple structure, small volume, light quality, convenience for control, and high safety. The aerodynamic high-performance aerofoil has a large real application value and provides novel thinking for aerofoil development.

There is provided a leading edge system for an aerospace vehicle. The leading edge system has at least one structural member. The leading edge system further has a plurality of removable modules removably attached to the at least one structural member. Each removable module has a hollow box substructure and at least one flange portion disposed along a first end of the hollow box substructure. The leading edge system further has a plurality of first attachment elements configured for attaching the at least one flange portion of the removable module to a first end portion of the at least one structural member. The leading edge system further has a plurality of second attachment elements configured for attaching a second end of the hollow box substructure opposite the flange portion, to a second end portion of the at least one structural member opposite the first end portion.

The invention relates to a variable camber wing trailing edge based on a compliant mechanism. The variable camber wing trailing edge based on the compliant mechanism is fixed at the rear end of a wingand provided with an elastic upper trailing edge skin and an elastic lower trailing edge skin. The trailing edge upper skin is fixed to a wing upper skin; and the trailing edge lower skin can retractinto a wing lower skin. An upper surface skin and a lower surface skin are connected by using a compliant driving mechanism; a connecting rod mechanism is additionally arranged between two points, with invariable distance, of the wing upper and lower skins before and after the deformation of a wing trailing edge to restrict the shape of the wing in the deformation process. Under the driving of adriving motor 1, the variable camber wing trailing edge based on the compliant mechanism enable a driving force to act upon the trailing edge skins by using the compliant mechanism so as to drive thetrailing edge upper skin to be bent downwards; and under the action of a driving motor 2, the trailing edge lower skin of the wing is driven to retract into the wing and is bent downwards. Through theadoption of two driving devices, the flexible trailing edge is guided to deform so as to change the aerodynamic performance of the wing.

The invention belongs to the field of aviation technology, in particular to a trailing edge and wing tip continuously variable camber structure and a decentralized driving control method thereof, theaircraft continuous camber structure of the present invention comprises SMA wires, for connecting the edges of two adjacent pterygoid joints on the same side of the hinge, and the SMA wire is connected with a heating power source and sealed with a plastic tube, A plastic tube is connected to the cooling compressor, An angle sensor is mounted on each hinge, SMA wires have shape memory function, Atlow temperature, that SMA wires can be elongate and deformed under the external force, and the shape of the fixed length can be kept. The SMA wire on one side is heated, and the SMA wires on the other side are cooled by the cool gas living in the house through a plastic tube, so that the whole wing-rib joint can be deflected to one side.

The invention provides a corporate aircraft engine upper placement and front swept wing duck type layout, which belongs to the aviation field, and relates to a high lift drag ratio airplane total layout scheme. The airplane comprises a fuselage, duck wings, wings, a vertical fin, an engine and a rear body edge strip; the wing leading edge inner segment sweeps back; the wing leading edge outer segment sweeps forward, and the engine is supported on an upper surface of a rear body edge strip between a wing trailing edge and an edge strip rudder by an engine, and in the top view projection, an engine axis crosses the corner of the wing leading edge inner segment and the wing leading edge outer segment, and an elevator is arranged at the wing trailing edge, and an inner side elevating auxiliary wing is arranged at the inner side of the wing trailing edge inner side, and an outside elevating auxiliary wing is arranged at the outer side of the wing trailing edge; an edge strip rudder is arranged at the rear edge of the rear body edge strip, and a rudder is arranged at the rear edge of the vertical fin. The wing trailing edge inner segment is connected with a rear body edge strip; the edge strip rudder at the rear body edge strip trailing edge is in cracking type, and is composed of an upper rudder surface and a lower rudder surface, and the shapes of the fuselage and the wing are in smooth transition. The invention has the advantages of high rising and landing lift coefficient and high trim lift drag ratio, light structure weight, long endurance and long voyage.

The invention relates to the technical field of aerospace, and discloses an active control method and device for aerofoil drag reduction. A plurality of piezoelectric devices are uniformly arranged at the trailing edge of an aerofoil in advance, and when an aircraft takes off or flies at a low speed, the piezoelectric devices are controlled to dent so as to enable concave pits to be formed in the corresponding position of the aerofoil, so that the separation of the boundary layer is postponed and the pressure drag is reduced; when the aircraft flies at a high speed, the piezoelectric devices are controlled to be bulged so as to enable bumps to be formed in the corresponding position, so that the drag is reduced; or the piezoelectric devices are controlled to be in the periodic oscillation operating mode, so that the wake zone of the aerofoil is enabled to form a periodic vortex structure so as to control flow separation. The active control method and device for aerofoil drag reduction have the advantages that the aircraft is enabled to actively switch the corresponding drag reduction mode in different flight regimes, and the flight efficiency is improved.

The invention provides a special spoon-shaped wing section with ultra-low Reynolds number, high lift-drag ratio and low speed. The thickness of the wing section before 60% of a chord is smaller and the thickness of the wing section after 60% of the chord is larger, so as to form a spoon shape geometry characteristic; the maximum relative thickness of the wing section before 60% of the chord is about 66% that of the wing section after 60% of the chord; the position of the maximum relative thickness of the wing section is at about 77% of the chord; the wing section has an area with reducing thickness at about 40% of the chord and the minimum relative thickness of the area of the wing section is about 35% of the maximum relative thickness of the wing section; the front part of the wing section is smaller in thickness and the rear part of the wing section is larger in thickness, so that the wing section has a good torque characteristic; under the Reynolds number of not greater than 10<4>, layer separation bubbles are small; the resistance of the wing section is greatly reduced, so that the wing section has a high lift-drag ratio and excellent pneumatic performance.

The invention discloses an air vehicle aerodynamic configuration with two trailing edge supporting wings. The supporting wings are installed below trailing edges of main wings at the two sides of a fuselage of an air vehicle. Wingtips of the supporting wings are connected to the middle parts of the trailing edges of the wings through connection sections. Wing roots of the supporting wings are connected to the fuselage through connection sections. The relative positions of the supporting wings and the main wings are reasonably designed. In a main wing cross section and a supporting wing cross section which are truncated by air flowing to a vertical plane in wingspans of the support wings upwardly during the flying process, the vertical distance between the centroids is a% of the chord length of the cross section of the main wing, wherein a is a constant value from 10 to 40; and meanwhile, the overlapping length of projections of the chord lines of the main wing cross section and the supporting wing cross section on a horizontal plane is b% of the chord length of the cross section of the main wing, wherein b is a constant value from 0 to 15. The aerodynamic configuration increases the total lift-drag ratio and achieves better aerodynamic performances, and also improves the rigidity of large-span-chord ratio wings and improves the total structural efficiency of the air vehicle.

An environmental protective coating (EPC) for protecting a surface subjected to high temperature environments of more than 3000 degree F. The coating includes a dense platelet lamellar microstructure with a self-sealing, compliant binder material for holding the platelets together. The platelets may be formed from materials that are resistant to high temperatures and impermeable, such as ceramics. The lamellar microstructure creates a tortuous path for oxygen to reach the surface. The binder material may have free internal volume to increase the strain capability between the platelets and absorb increased volume during operation. The binder may be formed from a material that is softer and has a lower temperature capability than the platelets to provide the system with the required compliance and sealing capability. The binder may have sufficient glass content and glass-forming content for initial and long-term sealing purposes.

An aircraft system includes a heat source and a passage near the heat source for carrying fluid having a cooling capacity to cool the heat source. The passage includes a catalyst that endothermically cracks the fluid to increase the cooling capacity.

An environmental protective coating (“EPC”) for protecting a surface subjected to high temperature environments of more than 3000 degree F. The coating includes a dense platelet lamellar microstructure with a self-sealing, compliant binder material for holding the platelets together. The platelets may be formed from materials that are resistant to high temperatures and impermeable, such as ceramics. The lamellar microstructure creates a tortuous path for oxygen to reach the surface. The binder material includes engineered free internal volume, which increases the elastic strain of the EPC. The binder is softer than the platelets, which in combination with its free volume increases pliability of the EPC. The binder may have sufficient glass content and glass-forming content for initial and long-term sealing purposes.