CN112249300B - Carbon fiber composite material airfoil leading edge structure - Google Patents
Carbon fiber composite material airfoil leading edge structure Download PDFInfo
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- CN112249300B CN112249300B CN202011136218.1A CN202011136218A CN112249300B CN 112249300 B CN112249300 B CN 112249300B CN 202011136218 A CN202011136218 A CN 202011136218A CN 112249300 B CN112249300 B CN 112249300B
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- leading edge
- box
- edge structure
- airfoil leading
- carbon fiber
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/28—Leading or trailing edges attached to primary structures, e.g. forming fixed slots
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/26—Construction, shape, or attachment of separate skins, e.g. panels
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
- B64U10/00—Type of UAV
- B64U10/25—Fixed-wing aircraft
Abstract
The invention relates to the technical field of unmanned aerial vehicles, in particular to a carbon fiber composite material airfoil leading edge structure. This carbon-fibre composite material airfoil leading edge structure, through the overall rigidity under the lightweight prerequisite of the layer structure that box-shaped frame and covering combine together, equipment joint frame adopts conducting material, electronic equipment passes through equipment joint frame and installs at airfoil leading edge structure, make carbon-fibre composite material airfoil leading edge structure have electronic equipment's support function, when using, equipment joint frame and other metal parts (like the wing girder) contact of unmanned aerial vehicle are electrically conductive, adopt flame spraying aluminium lamination to compare with the prior art in, volatility is littleer, and avoided the additional passive weight of flame spraying aluminium lamination and the risk of easily peeling off.
Description
Technical Field
The invention relates to the technical field of unmanned aerial vehicles, in particular to a carbon fiber composite material airfoil leading edge structure.
Background
The new generation aviation aircraft that uses high performance unmanned aerial vehicle as the representative is in order to subtract the needs that the weight increases the journey, adopt carbon-fibre composite material more, and its airfoil structure leading edge is because it bears the weight of and requires highly, the floor quantity is many, structural shape is irregular, compare the space with other structure positions of organism narrow and small relatively, and carbon-fibre composite material electric conductivity is unsatisfactory, so be not too easy as the carrier of electronic equipment installation, except satisfying the structural requirement, only as the passageway of threading in function usually, and the electronic equipment who carries uses an independent installing support to install usually. The mounting bracket of the electronic equipment has higher requirements on flatness, electric lap joint and the like, is made of metal materials, mechanically ensures form and position tolerances such as flatness and the like, and is mounted in a space-larger area such as a cabin body.
Disclosure of Invention
Technical problem to be solved
The invention aims to provide a carbon fiber composite material airfoil leading edge structure which has the function of an electronic equipment mounting bracket and solves at least one problem in the prior art.
(II) technical scheme
In order to achieve the above object, in a first aspect, the present invention provides a carbon fiber composite airfoil leading edge structure, including a skin as a main body of the airfoil leading edge structure, characterized in that: the inner part of the skin is provided with a plurality of box-shaped frames along the wingspan direction, two opposite side walls of the box-shaped frames are attached to the inner side of the skin, the end walls are positioned at the ends of the box-shaped frames and connected with the two opposite side walls, the two adjacent box-shaped frames form ribbed plates of a wing surface leading edge structure, the box-shaped frames form a plurality of ribbed plates arranged at intervals in the wingspan direction, the side walls of partial box-shaped frames are provided with equipment joint frames, the equipment joint frames are fixedly connected with the skin through bolt pairs, and the equipment joint frames are made of conductive materials.
Preferably, the equipment joint frame is further provided with a rivet hinge, the rivet hinge penetrates through the equipment joint frame, the side wall of the box-shaped frame and the skin, the rivet hinge and the bolt pair are arranged at intervals, and the axes of the rivet hinge and the bolt pair are parallel.
Preferably, the distance between the axis of the rivet hinge and the axis of the bolt pair is 3.5-5.5 times of the major diameter of the thread of the bolt pair.
Preferably, 1-4 layers of unidirectional prepreg are laid between two adjacent end walls of two adjacent box-shaped frames.
Preferably, the box-shaped frame adopts a quasi-isotropic layer, and the end wall of the box-shaped frame is formed by continuously laying layers of the side wall after the layers are bent.
Preferably, the carbon fiber composite airfoil leading edge structure has a plurality of equipment joint frames distributed on the respective sidewalls.
Preferably, the skin is further provided with a mounting edge, the mounting edge is laid on the inner surface of the side wall of the box-shaped frame corresponding to the edge of the skin, the innermost layer of the mounting edge is made of fabric, the rest of the laying layers are made of 0-degree unidirectional prepreg and the fabric alternately, and the number of adjacent 0-degree unidirectional prepreg layers is not more than three.
Preferably, the equipment joint frame has an extension section extending towards the edge of the skin, the extension section being in internal abutment with the mounting edge.
Preferably, the skin is symmetrically and symmetrically layered relative to the middle plane, 1-3 layers of plain woven fabric prepreg are layered on the inner and outer surfaces, a non-woven fabric prepreg is layered in the middle, and the proportion of 0-degree layering is 45-60%.
Preferably, the equipment connection frame is provided with a threaded connection port.
(III) advantageous effects
The technical scheme of the invention has the following advantages:
1. according to the carbon fiber composite material airfoil leading edge structure provided by the invention, the overall rigidity under the premise of light weight is ensured through the layering structure combining the box-shaped frame and the skin, the equipment joint frame is made of the conductive material, the electronic equipment is arranged on the airfoil leading edge structure through the equipment joint frame, so that the carbon fiber composite material airfoil leading edge structure has the function of a support of the electronic equipment, when the carbon fiber composite material airfoil leading edge structure is used, the equipment joint frame is in contact with other metal parts (such as a wing main beam) of an unmanned aerial vehicle for conducting, and compared with the aluminum layer sprayed by flame in the prior art, the volatility is smaller, and the additional negative weight and the risk of easy peeling of the aluminum layer sprayed by flame are avoided.
2. According to the other carbon fiber composite material airfoil leading edge structure, the plurality of equipment joint frames with smaller volumes are distributed at intervals, so that the design and the processing are convenient, the installation precision of equipment can be ensured by ensuring the form and position precision of the installed plane, the production cost is greatly reduced, and the equipment is connected into a whole by means of the self rigidity of the equipment after the equipment is installed.
Drawings
The drawings of the present invention are provided for illustrative purposes only, and the proportion and the number of the components in the drawings do not necessarily correspond to those of an actual product.
FIG. 1 is a schematic illustration of a carbon fiber composite airfoil leading edge configuration in accordance with an embodiment of the present invention;
FIG. 2 is another angular schematic view of the carbon fiber composite airfoil leading edge configuration of FIG. 1;
FIG. 3 is a schematic sectional view A-A of FIG. 2;
FIG. 4 is a schematic sectional view taken along line B-B of FIG. 2;
fig. 5 is an enlarged schematic view of the portion C in fig. 4.
In the figure: 1: covering a skin; 2: a box-shaped frame; 21: a side wall; 22: an end wall; 3: an equipment joint frame; 31: an extension section; 32: a threaded connector; 4: a bolt pair; 5: hinging and supporting a rivet; 6: and (7) installing edges.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention.
In general, in the field of aircraft design, an airfoil refers to a structure of a wing or a tail fin and a component thereof, and the airfoil mainly comprises a skin and a skeleton structure such as a spar and a rib in construction. Wherein the skin forms a streamlined wing or empennage outer surface. The airfoil leading edge refers to the portion of the wing or empennage forward of the spar in the span-wise direction indicated by the line connecting the two tips.
Referring to fig. 1 to 4, an airfoil leading edge structure made of a carbon fiber composite material provided by an embodiment of the present invention includes a skin 1 as a main body of the airfoil leading edge structure, a plurality of box frames 2 are disposed inside the skin 1 along a span direction, an outer shape of each box frame 2 matches an inner side of the skin 1, two opposite side walls 21 of each box frame 2 are attached to the inner side of the skin 1, end walls 22 are disposed at ends of the box frames 2 and connect the two opposite side walls, two adjacent box frames 2, two adjacent end walls 22 of each box frame 2 form a rib plate of the airfoil leading edge structure, the plurality of box frames 2 form a plurality of rib plates disposed at intervals in the span direction, an equipment joint frame 3 is disposed on the side walls 21 of a part of the box frames 2, the equipment joint frame 3 is fixedly connected to the skin 1 by a bolt pair 4, the equipment joint frame 3 is made of a conductive material, for example, metal such as iron, steel, alloy such as copper alloy, aluminum alloy, etc., alloy, etc, Conductive plastics, conductive rubber, conductive fiber fabric and other composite polymer conductive materials.
The carbon fiber composite material airfoil leading edge structure ensures the integral rigidity on the premise of light weight through a layer spreading structure combining a box-shaped frame and a skin, the equipment joint frame 3 is made of a conductive material, and the electronic equipment is installed on the airfoil leading edge structure through the equipment joint frame 3, so that the carbon fiber composite material airfoil leading edge structure has the support function of the electronic equipment. In some embodiments, the contact resistance over a 200mm length can be up to 0.17m Ω, well below 6-16 m Ω of the flame sprayed aluminum layer.
In some preferred embodiments, referring to fig. 3 and 4, the equipment joint frame 3 is further provided with a rivet hinge 5, the rivet hinge 5 penetrates through the equipment joint frame 3, the side wall 21 of the box-shaped frame and the skin 1, and the rivet hinge 5 can rotate relative to the equipment joint frame 3, the side wall 21 of the box-shaped frame and the skin 1 to form a hinge point. The rivet hinge 5 and the bolt pair 4 are arranged at intervals, and the axes of the rivet hinge and the bolt pair are parallel. In the embodiment, the equipment joint frame 3 is connected with other parts by adopting a rivet hinge 5 and a bolt pair 4, has certain rotational freedom degree on the premise of ensuring the connection strength, can realize coordinated deformation with equipment under working conditions such as vibration and the like, and meets the requirements of mechanical environment. Preferably, the axial distance between the rivet hinge support 5 and the bolt pair 4 is 3.5-5.5 times of the major diameter of the thread of the bolt pair 4.
In some preferred embodiments, referring to fig. 1 and 2, a plurality of small-sized device connector frames 3 are distributed at intervals, and the device connector frames 3 are arranged on the corresponding side walls 21 according to the electronic device installation requirements, for example, three device connector frames 3 are respectively arranged on two opposite side walls 21 in one box-shaped frame 2, three device connector frames 3 on each side wall 21 are arranged at intervals, and two device connector frames on two side walls 21 are arranged oppositely. The equipment joint frame 3 with a small size is adopted in the implementation mode, so that the design and the processing are convenient, the installation precision of the equipment can be ensured by ensuring the shape and position precision of the plane after the installation, the production cost is greatly reduced, and the equipment joint frame is connected into a whole by means of the rigidity of the equipment after the equipment is installed.
In some preferred embodiments, two adjacent box frames 2, with layers of unidirectional prepreg 1-4 (not shown) laid between two adjacent end walls 22, can increase unidirectional stiffness and strength.
In some preferred embodiments, the skin 1 is uniformly and symmetrically laid on the middle surface, 1-3 layers of plain woven fabric prepregs are laid on the inner surface and the outer surface, and a non-woven fabric prepreg is laid on the middle surface, wherein the proportion of 0-degree laying is 45-60%, and the skin has high strength. If the reinforcing rib plate is combined with the implementation mode that 1-4 layers of unidirectional prepreg are laid in the rib plate, the reinforcing rib plate can avoid curing deformation to the maximum extent on the premise of improving the strength, and the size stability of the support is ensured.
In some preferred embodiments, the box frame 2 is formed from a quasi-isotropic laminate, the end walls 22 of which are formed by folding and then continuously laying the laminate on the side walls 21.
In some preferred embodiments, as shown in fig. 1, 3, 4 and 5, the edge inside the covering skin is provided with a mounting edge 6, and the layer of the mounting edge is positioned inside the layer of the side wall 21 of the box-shaped frame when the layers are laid, namely, the covering layer, the layer of the side wall of the box-shaped frame and the layer of the mounting edge are arranged on the layer structure from outside to inside. The mounting edge 6 is arranged for connection to other parts of the airfoil for increased structural strength. Preferably, the innermost layer (the layer far away from the skin 1) of the mounting edge 6 is made of fabric, the rest of the layers are made of unidirectional prepreg and fabric alternately, and the number of the adjacent 0-degree unidirectional prepreg layers is not more than three. Preferably, the equipment joint frame 3 has an extension 31 extending toward the edge of the skin 1, the extension 31 being in-layer butt-jointed with the mounting edge 6.
In some preferred embodiments, the equipment connection block is provided with a threaded connection port 32 for connection with an electronic device.
In one particular embodiment, and referring to FIGS. 1-5, a carbon fiber composite airfoil leading edge structure is 1500mm in length and approximately isosceles triangular in cross-section with a base 190mm in length and a height 230 mm.
Four box-shaped frames 2 are arranged on the inner side of the skin 1, two surfaces of the two box-shaped frames 2 positioned on the outer sides of two ends in the wingspan direction are open, and one surface of each of the two box-shaped frames 2 positioned in the middle is open. In the middle two box-shaped frames 2, two equipment joint frames 3 are arranged on each side wall 21 at intervals, and the total number of the equipment joint frames 3 is eight, and each equipment joint frame 3 is provided with an extension section 31. Each equipment joint frame 3 is connected with the skin through a phi 5 rivet hinge 5 and an M6 bolt pair 4, and the integral flatness of the equipment mounting surface formed after assembly does not exceed 0.15. The axial distance between the rivet hinge 5 and the bolt pair 4 is 30 mm.
Layers (not shown) of unidirectional prepreg 2 are laid between adjacent end walls 22 of two adjacent box frames 2 to form ribs together. The box-shaped frame 2 adopts quasi-isotropic ply, [ (0,90)/(45, -45)/(0,90) ] s, the plies of the side wall 21 and the end wall 22 are continuous, and the plies of the end wall 22 are continuously paved after being bent by the plies of the side wall 21. Preferably, the ratio between the thickness of the end wall 22 and the thickness of the side wall 21 is 1: 2, the side walls 21 are 1.5mm thick and the end walls 22 are 3mm thick.
The skin 1 is of an integral structure, and the layers are as follows: the inner and outer surface layers adopt 1-3 layers of plain woven fabric prepreg, the middle layer adopts non-woven fabric prepreg, the proportion of 0-degree layer is 45-60%, and the specific layer is [0 ]2/902/02/90]s。
The mounting edge 6 is an additional layer of the inner surface, the width is 75mm, the thickness is 0.75mm, and the specific layer is as follows: [(0,90)/03/(0,90)]And the extension section of the equipment joint frame 2 is in-layer butt joint.
The equipment joint frame 3 is a box-shaped structure with only one open face, and the thickness of the extension 31 is 0.75 mm.
It should be noted that, referring to fig. 1, the box-shaped frame is a box-shaped frame structure having a cavity and formed by a side wall and an end wall, and has at least one opening communicated with the cavity. When the box frame is at the extreme edge of the mask, two openings may be provided (one end wall less than the box frame in the middle).
It should be further noted that the device connector frame is a frame structure with an open cavity formed by several surrounding surfaces, and can accommodate electronic devices or a part of the electronic devices to realize connection and fixation of the electronic devices.
It should be noted that the bolt pair 4 at least comprises a bolt and a nut matched with the bolt, and mainly plays a role of reinforcing the joint frame 3 of the fixing equipment in the invention.
It should be noted that fig. 1-5 are schematic diagrams provided for the structures of the leading edge portions of the airfoil, and after the curing and forming, the adjacent layers of different portions are integrally formed, for example, between the mounting edge 6 and the extension 31, between the mounting edge 6 and the extension and the box frame 2, between the box frame 2 and the skin 1, and between two adjacent end walls 22.
Finally, it should be noted that: the above examples are only intended to illustrate the technical solution of the present invention, but not to limit it; although the present invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: each embodiment does not include only one independent technical solution, and in the case of no conflict between the solutions, the technical features mentioned in the respective embodiments can be combined in any way to form other embodiments which can be understood by those skilled in the art.
Furthermore, modifications may be made to the technical solutions described in the foregoing embodiments, or equivalents may be substituted for some of the technical features thereof, without departing from the scope of the present invention, and the essence of the corresponding technical solutions does not depart from the spirit and scope of the technical solutions of the embodiments of the present invention.
Claims (7)
1. A carbon fiber composite airfoil leading edge structure comprising a skin as an airfoil leading edge structure body, characterized in that: the interior of the skin is provided with a plurality of box-shaped frames along the wingspan direction, two opposite side walls of the box-shaped frames are attached to the inner side of the skin, end walls are positioned at the ends of the box-shaped frames and connected with the two opposite side walls, two adjacent box-shaped frames form ribbed plates of an airfoil leading edge structure, the box-shaped frames form a plurality of ribbed plates arranged at intervals in the wingspan direction, equipment joint frames are arranged on part of the side walls of the box-shaped frames and connected and fixed with the skin through bolt pairs, the equipment joint frames are made of conductive materials, the carbon fiber composite material airfoil leading edge structure is provided with a plurality of equipment joint frames, the equipment joint frames are distributed on the corresponding side walls, the equipment joint frames are of a structure with only one open surface, mounting edges are further arranged in the skin, and are laid on the inner surfaces of the side walls of the box-shaped frames corresponding to the edges of the box-shaped frames, the innermost layer of the mounting edge is made of fabric, the rest layers are made of 0-degree unidirectional prepreg and fabric alternately, and the number of adjacent 0-degree unidirectional prepreg layers is not more than three;
the equipment joint frame is provided with an extension section extending towards the edge of the skin, and the extension section is in butt joint with the installation edge layer.
2. The carbon fiber composite airfoil leading edge structure of claim 1, wherein: the equipment joint frame is further provided with a rivet hinge which penetrates through the equipment joint frame, the side wall of the box-shaped frame and the skin, the rivet hinge and the bolt pair are arranged at intervals, and the axes of the rivet hinge and the bolt pair are parallel.
3. The carbon fiber composite airfoil leading edge structure of claim 2, wherein: the distance between the rivet hinge support and the axis of the bolt pair is 3.5-5.5 times of the major diameter of the thread of the bolt pair.
4. The carbon fiber composite airfoil leading edge structure of claim 1, wherein: and 1-4 layers of unidirectional prepreg are laid between two adjacent end walls of the two adjacent box-shaped frames.
5. The carbon fiber composite airfoil leading edge structure of claim 1, wherein: the box-shaped frame adopts a quasi-isotropic layer, and the end wall of the box-shaped frame is formed by continuously laying and covering the layer of the side wall after being bent.
6. The carbon fiber composite airfoil leading edge structure of claim 1, wherein: the skin adopts the balanced symmetrical shop of relative mid-plane to spread the layer, and interior external surface shop lays and adopts 1~3 layers plain weave preimpregnation material, and the middle shop lays and adopts no latitude cloth preimpregnation material, and the proportion of 0 shop laying is between 45~ 60%.
7. The carbon fiber composite airfoil leading edge structure of claim 1, wherein: the equipment joint frame is provided with a threaded connector.
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| CN202011136218.1A CN112249300B (en) | 2020-10-22 | 2020-10-22 | Carbon fiber composite material airfoil leading edge structure |
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| CN202011136218.1A CN112249300B (en) | 2020-10-22 | 2020-10-22 | Carbon fiber composite material airfoil leading edge structure |
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| CN112249300B true CN112249300B (en) | 2022-02-15 |
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| US8968437B2 (en) * | 2012-05-02 | 2015-03-03 | Michael J Kline | Jet engine with deflector |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2011051517A (en) * | 2009-09-03 | 2011-03-17 | Mitsubishi Heavy Ind Ltd | Aircraft wing |
| EP3006189A1 (en) * | 2014-10-09 | 2016-04-13 | The Boeing Company | Composite structures having composite-to-metal joints and methods for making the same |
| CN106114819A (en) * | 2016-07-22 | 2016-11-16 | 中国航空工业集团公司西安飞机设计研究所 | A kind of composite airfoil structure |
| CN207157531U (en) * | 2017-08-23 | 2018-03-30 | 周袭明 | Flapping-wing aircraft fixed-wing and flapping-wing aircraft |
| CN108216570A (en) * | 2017-12-14 | 2018-06-29 | 中国航空工业集团公司成都飞机设计研究所 | A kind of high aspect ratio wing main plane structure |
| EP3428056A1 (en) * | 2017-07-13 | 2019-01-16 | Airbus Operations, S.L. | Box structural arrangenment for an aircraft and manufacturing method thereof |
| CN109502007A (en) * | 2017-09-15 | 2019-03-22 | 空中客车运营简化股份公司 | A kind of wing and the aircraft including this wing |
| CA3067469A1 (en) * | 2019-01-14 | 2020-07-14 | The Boeing Company | Aircraft wing composite ribs having electrical grounding paths |
Family Cites Families (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO1997034734A1 (en) * | 1996-03-22 | 1997-09-25 | The Boeing Company | Determinant wing assembly |
| FI118761B (en) * | 2006-02-09 | 2008-03-14 | Patria Aerostructures Oy | Wing for aircraft, mounting arrangement and intermediate support |
| BRPI0805129A2 (en) * | 2008-11-18 | 2010-08-17 | Jose Antonio Romano Espinosa | airfoil with hollow depressions on the surface |
| US8123167B2 (en) * | 2008-12-15 | 2012-02-28 | Embraer S.A. | Impact resistant aircraft leading edge structures and aircraft including the same |
| US9595910B2 (en) * | 2011-07-18 | 2017-03-14 | James L. Grimsley | Solar assembly and method of forming same |
| US9227718B2 (en) * | 2013-06-07 | 2016-01-05 | The Boeing Company | Lower joints between outboard wing boxes and center wing sections of aircraft wing assemblies |
| CN103434637A (en) * | 2013-08-25 | 2013-12-11 | 西北工业大学 | Novel aerofoil by utilizing magnus effect |
| CN103538715B (en) * | 2013-10-14 | 2016-06-08 | 航天特种材料及工艺技术研究所 | A kind of matrix material �� type ear blade terminal and overall curing molding method altogether thereof |
| EP2962840A1 (en) * | 2014-06-30 | 2016-01-06 | Airbus Operations, S.L. | A leading edge for an aircraft lifting surface and manufacturing method thereof |
| US10843416B2 (en) * | 2015-05-11 | 2020-11-24 | Gulfstream Aerospace Corporation | Composite reinforcement structures and aircraft assemblies comprising composite reinforcement structures |
| EP3374262B1 (en) * | 2015-11-09 | 2021-02-24 | Sikorsky Aircraft Corporation | Rotor blade structures |
| CN208181397U (en) * | 2018-01-30 | 2018-12-04 | 彩虹无人机科技有限公司 | A kind of leading edge of a wing lightweight bracket |
| US20200331616A1 (en) * | 2019-04-19 | 2020-10-22 | Goodrich Corporation | Bonded structural rib for heated aircraft leading edge |
-
2020
- 2020-10-22 CN CN202011136218.1A patent/CN112249300B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2011051517A (en) * | 2009-09-03 | 2011-03-17 | Mitsubishi Heavy Ind Ltd | Aircraft wing |
| EP3006189A1 (en) * | 2014-10-09 | 2016-04-13 | The Boeing Company | Composite structures having composite-to-metal joints and methods for making the same |
| CN106114819A (en) * | 2016-07-22 | 2016-11-16 | 中国航空工业集团公司西安飞机设计研究所 | A kind of composite airfoil structure |
| EP3428056A1 (en) * | 2017-07-13 | 2019-01-16 | Airbus Operations, S.L. | Box structural arrangenment for an aircraft and manufacturing method thereof |
| CN207157531U (en) * | 2017-08-23 | 2018-03-30 | 周袭明 | Flapping-wing aircraft fixed-wing and flapping-wing aircraft |
| CN109502007A (en) * | 2017-09-15 | 2019-03-22 | 空中客车运营简化股份公司 | A kind of wing and the aircraft including this wing |
| CN108216570A (en) * | 2017-12-14 | 2018-06-29 | 中国航空工业集团公司成都飞机设计研究所 | A kind of high aspect ratio wing main plane structure |
| CA3067469A1 (en) * | 2019-01-14 | 2020-07-14 | The Boeing Company | Aircraft wing composite ribs having electrical grounding paths |
Non-Patent Citations (1)
| Title |
|---|
| 先进复合材料在无人机上的应用及关键技术;杜龙;《教练机》;20170615;第10-17页 * |
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