CN114248905B - Composite wing spar and forming method thereof - Google Patents
Composite wing spar and forming method thereof Download PDFInfo
- Publication number
- CN114248905B CN114248905B CN202111392164.XA CN202111392164A CN114248905B CN 114248905 B CN114248905 B CN 114248905B CN 202111392164 A CN202111392164 A CN 202111392164A CN 114248905 B CN114248905 B CN 114248905B
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- composite material
- wing spar
- integral beam
- composite
- spar
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- 239000002131 composite material Substances 0.000 title claims abstract description 72
- 238000000034 method Methods 0.000 title claims abstract description 19
- 230000007704 transition Effects 0.000 claims abstract description 17
- 238000004519 manufacturing process Methods 0.000 claims description 11
- 238000010438 heat treatment Methods 0.000 claims description 7
- 238000004140 cleaning Methods 0.000 claims description 5
- 238000001816 cooling Methods 0.000 claims description 5
- 238000007711 solidification Methods 0.000 claims description 4
- 230000008023 solidification Effects 0.000 claims description 4
- 239000003795 chemical substances by application Substances 0.000 claims description 3
- 239000000428 dust Substances 0.000 claims description 3
- 239000012535 impurity Substances 0.000 claims description 3
- 239000012945 sealing adhesive Substances 0.000 claims description 3
- 239000003921 oil Substances 0.000 claims description 2
- 239000000446 fuel Substances 0.000 abstract description 3
- 238000001723 curing Methods 0.000 description 6
- 238000005265 energy consumption Methods 0.000 description 4
- 239000011248 coating agent Substances 0.000 description 2
- 238000000576 coating method Methods 0.000 description 2
- 238000003754 machining Methods 0.000 description 2
- 229920000049 Carbon (fiber) Polymers 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000004917 carbon fiber Substances 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000013007 heat curing Methods 0.000 description 1
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000000465 moulding Methods 0.000 description 1
- 230000003014 reinforcing effect Effects 0.000 description 1
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/18—Spars; Ribs; Stringers
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/30—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
- B29C70/34—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation
- B29C70/342—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation using isostatic pressure
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/54—Component parts, details or accessories; Auxiliary operations, e.g. feeding or storage of prepregs or SMC after impregnation or during ageing
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/30—Vehicles, e.g. ships or aircraft, or body parts thereof
- B29L2031/3076—Aircrafts
- B29L2031/3085—Wings
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Chemical & Material Sciences (AREA)
- Composite Materials (AREA)
- Aviation & Aerospace Engineering (AREA)
- Moulding By Coating Moulds (AREA)
Abstract
The invention discloses a wing spar made of composite materials and a forming method thereof, wherein the wing spar is an integrally formed integral beam, a plurality of cavities are arranged along the length direction of the integral beam, and the inner wall surfaces of the cavities are web surfaces of the integral beam; the upper surface of the integral beam is provided with an upper edge strip surface except the cavity, and the upper edge strip surface is provided with an inclined surface and is used for assembling an upper skin; a lower flanging for assembling the lower skin is arranged on the surface opposite to the upper edge strip surface; the web surface is in transition connection with the upper edge strip surface through a round angle; the two adjacent web surfaces are in transition connection through a round angle; the side surface of the integral beam is in transition connection with the lower flanging through a round angle. The wing spar of the composite material is an integrated integral spar structure, so that the number of components of the spar is reduced, and the integrity and reliability of the spar structure are improved; compared with the traditional spar structure, the integral beam has the advantages that the number of fasteners is reduced, the weight of the spar is reduced, the cost is reduced, and the fuel economy and the maneuverability of an airplane are improved.
Description
Technical Field
The invention relates to the technical field of aviation equipment, in particular to a composite wing spar and a forming method thereof.
Background
Wings are one of the important components of an aircraft and are mounted on the aircraft fuselage. The main function of the wing is to generate lift and form good stability and maneuverability together with the tail wing. As shown in fig. 1, the main structural components of the conventional aircraft wing are: upper skin, lower skin, spar, rib, stringer, etc.; the wing spar is a main bearing part of the wing and is a core part for transmitting the wing load.
At present, the spar structure of the traditional carbon fiber composite wing needs to be assembled with structural members such as various ribs, stringers, reinforcing ribs and the like to bear, so that the number of parts is large, the number of dies is large, and the spar of the structure mainly has the following defects: (1) The number of the molds required for manufacturing is large, so that the cost of the molds is high; (2) The number of parts is large, the assembly tool is complex, and the assembly tool has high cost; (3) The product needs to enter a curing furnace for curing for many times, and the equipment and energy consumption are high; (4) The subsequent assembly needs to be performed respectively, the assembly time is more, the production efficiency is low, and the assembly cost is high.
Disclosure of Invention
Aiming at the problems of high cost and the like caused by more mould demands, more spar parts and high production energy consumption and the problems of low production efficiency and the like caused by complex assembly and time consumption of the traditional wing spar structure, the invention provides a novel composite wing spar and a forming method thereof.
The invention is realized by the following technical scheme:
The wing spar of the composite material is characterized in that the wing spar is an integrally formed integral beam, a plurality of cavities are formed along the length direction of the integral beam, and the inner wall surfaces of the cavities are web surfaces (a plurality of web surfaces) of the integral beam; the upper surface of the integral beam is provided with an upper edge strip surface (namely the upper edge strip surface of the integral beam) except the cavity, and the upper edge strip surface is provided with an inclined surface and is used for assembling an upper skin; a lower flanging for assembling the lower skin is arranged on the surface, opposite to the upper edge strip surface, of the integral beam (namely, the lower surface of the integral beam is provided with the lower flanging for assembling the lower skin); the web plate surface is in transition connection with the upper edge strip surface through a round angle; the two adjacent web surfaces are in transition connection through a round angle; the side surface of the integral beam is in transition connection with the lower flanging through a round angle. Specifically, the cavity is rectangular, and the inner wall surface of the cavity is a web surface, so that a plurality of web surfaces exist in one cavity.
Specifically, according to the composite wing spar, the bearing structure of the traditional wing spar comprises the wing ribs, the stringers, the spars and the like which are designed into an integrated spar structure (integral beam), so that the number of parts of the spar is reduced, and the integrity and the reliability of the spar structure are improved; the integral beam provided by the invention reduces the number of fasteners, reduces the structural weight, reduces the manufacturing cost and improves the fuel economy and maneuverability of an aircraft. The composite wing spar (integral beam) has less demand on the forming die, and the die cost is reduced. The composite wing spar (integral beam) has the advantages of less parts, simple assembly, low production efficiency and low assembly cost. The integral beam does not need to be heated and solidified for many times, and has low production energy consumption and lower cost.
The wing spar (integral beam) of the composite material mainly comprises an upper edge strip surface, a web surface, a downward flanging and a plurality of typical areas on the side surface of the integral beam. The section of the integral beam can be regarded as being similar to a hat-shaped omega-shaped structure, the upper panel of the structure is an assembly area bonded with the upper skin (namely, the upper edge strip surface is provided with an inclined surface and is used for accelerating the gas flow rate and providing lift force), and the upper panel and the upper skin assembly area have the same curvature; the lower part of the structure is an assembly area (lower flanging) which is adhered with the lower skin, and the curvature of the lower flanging is the same as that of the lower skin assembly area.
Further, the composite wing spar: the width of the downward flanging assembled with the lower skin is not less than 20mm.
Further, the composite wing spar: the web plate surface and the upper edge strip surface are in transition connection by adopting a round angle with the diameter not smaller than 2.5 mm; the web surfaces are connected in a transition manner by adopting a round angle with the diameter not smaller than 2.5 mm; and the side surface of the integral beam is in transition connection with the lower flanging by adopting a round angle with the diameter not smaller than 2.5 mm.
A method of forming a composite wing spar, the method comprising the steps of:
s1, manufacturing a forming die consistent with the integral beam structure;
S2, cleaning the forming die;
S3, paving a composite material on the forming die according to the loaded characteristic of the wing; the paving process of the composite material comprises the following steps: taking the spanwise direction as the 0-degree direction and the chordwise direction as the 90-degree direction; composite lay-up direction 0 °/±45°/90 °: the composite material layering proportion of the upper edge strip surface (2) and the lower flanging (3) is respectively 60%/20%/20%, and the composite material layering proportion of the web surface (1) is respectively 30%/40%/30%;
S4, after the step S3 of paving the composite material is completed, the composite material is integrally sealed, and then vacuumizing is carried out;
S5, vacuumizing and heating to cure the composite material;
And S6, cooling after solidification is completed, and demoulding to obtain the wing spar made of the composite material.
Further, the method for forming the wing spar of the composite material comprises the following steps: and S1, machining a forming die consistent with the integral beam structure according to the size of the integral beam to be formed.
Further, the method for forming the wing spar of the composite material comprises the following steps: and S2, cleaning the forming die, removing oil stains, dust and impurities on the die, and then coating a release agent.
Further, the method for forming the wing spar of the composite material comprises the following steps: and S3, paving the composite material with the thickness not smaller than 1mm.
Further, the method for forming the wing spar of the composite material comprises the following steps: and S4, after the composite material in the step S3 is paved, the composite material is integrally coated by a vacuum bag, and is sealed by a sealing adhesive tape to ensure air tightness, and then the vacuum is pumped until the vacuum degree is minus 0.98bar plus or minus 0.01bar.
Further, the method for forming the wing spar of the composite material comprises the following steps: and S5, placing the molding die into an autoclave after vacuumizing is finished, heating to 118-122 ℃ at a speed of 1-3 ℃/min for curing, and continuously maintaining the temperature for curing for 1-2 hours after heating is finished.
Further, the method for forming the wing spar of the composite material comprises the following steps: and S6, cooling to 55-65 ℃ after solidification, and demolding to obtain the wing spar made of the composite material, namely the integral beam.
The invention has the beneficial effects that:
(1) The composite wing spar manufactured by the invention has the advantages that the bearing structure of the traditional wing spar comprises the wing ribs, the stringers, the spars and the like, which are designed into an integrated integral spar structure, the structure is simple, the number of components of the spar is reduced, and the integrity and the reliability of the spar structure are improved; compared with the traditional wing spar structure, the integral beam provided by the invention has the advantages that the number of fasteners is greatly reduced, the weight of the spar is reduced, the manufacturing cost is reduced, and the fuel economy and the maneuverability of an airplane are improved.
(2) The number of die tools for manufacturing the wing spar (integral beam) made of the composite material can be greatly reduced, and the die cost can be greatly reduced; and the number of the assembly tools can be greatly reduced, and the cost of the assembly tools can be greatly reduced.
(3) The composite wing spar provided by the invention has the advantages that the required heat curing times are small, and the energy consumption is low; the assembly process and working hour are few, and the assembly cost is low.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings that are needed in the description of the embodiments will be briefly described below, it being obvious that the drawings in the following description are only some embodiments of the present invention, and that other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a schematic structural view of a conventional wing;
FIG. 2 is a schematic structural view of a wing spar of the present invention;
FIG. 3 is a top view of a wing spar of the present invention;
FIG. 4 is a cross-sectional view taken along the direction A-A in FIG. 3;
FIG. 5 is a cross-sectional view taken along the direction B-B in FIG. 3;
FIG. 6 is a cross-sectional view taken along the direction C-C in FIG. 3;
FIG. 7 is a cross-sectional view taken along the direction D-D in FIG. 3;
fig. 8 is an exploded view of a wing made in accordance with the present invention.
In the figure: 1 web surface, 2 upper edge strip surface, 3 lower flanging, 4 side surfaces, 5 upper skin, 6 lower skin, 7 spar, 8 rib and 9 end rib.
Detailed Description
The following description of the embodiments of the present invention will be made clearly and fully with reference to the accompanying drawings, in which it is evident that the embodiments described are only some, but not all embodiments of the invention. The following description of at least one exemplary embodiment is merely exemplary in nature and is in no way intended to limit the invention, its application, or uses. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
In the description of the present invention, it should be understood that the terms "upper," "lower," "left," "right," "top," "bottom," and the like indicate orientations or positional relationships, merely to facilitate describing the present invention and simplify the description, and do not indicate or imply that the devices or elements being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus should not be construed as limiting the invention. Furthermore, the terms "first," "second," and the like, are used for descriptive purposes only and are not to be construed as indicating or implying a relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defining "a first" or "a second" may include one or more of the feature, either explicitly or implicitly. Moreover, the terms "first," "second," and the like, are used for distinguishing between similar objects and not necessarily for describing a particular sequential or chronological order. It is to be understood that the data so used may be interchanged where appropriate such that the embodiments of the invention described herein may be implemented in sequences other than those illustrated or otherwise described herein.
Example 1
As shown in fig. 2-7, the wing spar made of composite material is an integrally formed integral beam, and a plurality of cavities are arranged at equal intervals along the length direction of the integral beam, the inner wall surfaces of the cavities are web surfaces 1 of the integral beam, the area of the upper surface of the integral beam except the cavities is an upper edge strip surface 2 of the integral beam, the upper edge strip surface 2 is provided with an inclined surface, and the upper edge strip surface 2 is used for being bonded and assembled with an upper skin 5, so that the main function is to generate lifting force required by taking off an aircraft; a lower flange 3 for assembling the lower skin 6 is arranged on the surface of the integral beam opposite to the upper edge strip surface 2 (namely, the lower surface of the integral beam is provided with the lower flange 3 for assembling the lower skin 6, and the width of the lower flange 3 is 25 mm); the web surface 1 is in transitional connection with the upper edge strip surface 2 through a round angle with the diameter of 2.5 mm; the two adjacent web surfaces 1 are in transition connection through a round angle with the diameter of 2.5mm, and the side surface 4 of the integral beam is in transition connection with the lower flanging 3 through a round angle with the diameter of 2.5 mm.
The method for forming the wing spar of the composite material comprises the following steps of:
S1, machining a forming die consistent with the integral beam structure according to the size of the integral beam to be formed;
s2, cleaning the forming die, removing greasy dirt, dust and impurities on the die, and then coating a release agent;
S3, paving a composite material with the thickness not less than 1 mm on the forming die according to the loaded characteristic of the wing; the paving process of the composite material comprises the following steps: the spanwise direction is taken as the 0-degree direction, and the chordwise direction is taken as the 90-degree direction (as shown in fig. 3); composite lay-up direction 0 °/±45°/90 °: the composite material layering ratio of the upper edge strip surface 2 and the lower flanging 3 is 60%/20%/20%; the composite material layering ratio of the web surface 1 in the direction of 0 degree/45 degree/90 degree is 30 percent/40 percent/30 percent respectively;
S4, after the step S3 of paving the composite material is completed, the composite material is integrally coated by a vacuum bag, and is sealed by a sealing adhesive tape to ensure air tightness, and then the vacuum bag is vacuumized through a vacuum joint until the vacuum degree is-0.98 bar;
s5, after the vacuumizing is finished, placing the forming die in an autoclave, heating to 120 ℃ at a speed of 1 ℃/min for curing, and continuously maintaining 120 ℃ for curing the paved composite material for 1.5 hours after the heating is finished;
S6, naturally cooling to 60 ℃ after solidification is completed, and demolding to obtain the wing spar made of the composite material, namely the integral beam. In order to ensure that the product can be smoothly demoulded, the side surface 4 of the integral beam adopts a demould slope of more than or equal to 1 DEG, and the width of the lower flanging 3 assembled with the lower skin is set to be 25mm.
The upper and lower surfaces of the wing spar of the composite material prepared by the invention are respectively adhered with an upper skin 5 and a lower skin 6, and then the whole wing of the composite material is obtained after the end ribs 9 are assembled at the two ends (as shown in figure 8).
The above-described preferred embodiments of the present invention are only for illustrating the present invention, and are not to be construed as limiting the present invention. Obvious changes and modifications of the invention, which are introduced by the technical solution of the present invention, are still within the scope of the present invention.
Claims (9)
1. The wing spar of the composite material is characterized in that the wing spar is an integrally formed integral beam, a plurality of cavities are formed along the length direction of the integral beam, and the inner wall surfaces of the cavities are web surfaces (1) of the integral beam; the upper surface of the integral beam is provided with an upper edge strip surface (2) except the cavity, and the upper edge strip surface (2) is provided with an inclined surface and is used for assembling an upper skin; a lower flanging (3) for assembling a lower skin is arranged on the surface, opposite to the upper edge strip surface (2), of the integral beam; the web plate surface (1) is in transition connection with the upper edge strip surface (2) through a round angle; the two adjacent web surfaces (1) are connected through a fillet transition; the side surface (4) of the integral beam is in transition connection with the lower flanging (3) through a round angle;
the method for forming the wing spar of the composite material comprises the following steps:
s1, manufacturing a forming die consistent with the integral beam structure;
S2, cleaning the forming die;
S3, paving a composite material on the forming die according to the loaded characteristic of the wing; the paving process of the composite material comprises the following steps: taking the spanwise direction as the 0-degree direction and the chordwise direction as the 90-degree direction; composite lay-up direction 0 °/±45°/90 °: the composite material layering proportion of the upper edge strip surface (2) and the lower flanging (3) is respectively 60%/20%/20%, and the composite material layering proportion of the web surface (1) is respectively 30%/40%/30%;
S4, after the step S3 of paving the composite material is completed, the composite material is integrally sealed, and then vacuumizing is carried out;
S5, vacuumizing and heating to cure the composite material;
And S6, cooling after solidification is completed, and demoulding to obtain the wing spar made of the composite material.
2. A composite wing spar according to claim 1, wherein the width of the turndown (3) fitted with the lower skin is not less than 20mm.
3. A composite wing spar according to claim 1, wherein the web surface (1) and the upper edge strip surface (2) are connected by a fillet transition having a diameter of not less than 2.5 mm; the web surfaces (1) are connected in a transition manner by adopting a round angle with the diameter not smaller than 2.5 mm; and the side surface (4) of the integral beam is in transition connection with the lower flanging (3) by adopting a round angle with the diameter not smaller than 2.5 mm.
4. A composite wing spar according to claim 1, wherein step S1, the integral beam being formed according to the dimensions of the integral beam being formed, is machined to form a forming mould conforming to the integral beam structure.
5. A composite wing spar according to claim 1, wherein step S2, cleaning the mould, removing oil, dust and impurities from the mould, and then applying a release agent.
6. A composite wing spar according to claim 1, wherein the lay-up thickness of the composite material in step S3 is not less than 1mm.
7. The wing spar of the composite material according to claim 1, wherein after the composite material of the step S4 and the step S3 is laid, the composite material is integrally covered by a vacuum bag, and is sealed by a sealing adhesive tape to ensure air tightness, and then the wing spar is vacuumized until the vacuum degree is-0.98 bar+/-0.01 bar.
8. The wing spar of claim 1, wherein after the step S5, the forming mold is placed in an autoclave after the vacuumizing is completed, and the forming mold is heated to 118-122 ℃ at a speed of 1-3 ℃/min for curing, and the curing is continued for 1-2 hours after the heating is completed.
9. The composite wing spar according to claim 1, wherein the composite wing spar is obtained by cooling to 55-65 ℃ after curing in step S6 and then demoulding.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111392164.XA CN114248905B (en) | 2021-11-23 | 2021-11-23 | Composite wing spar and forming method thereof |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111392164.XA CN114248905B (en) | 2021-11-23 | 2021-11-23 | Composite wing spar and forming method thereof |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN114248905A CN114248905A (en) | 2022-03-29 |
| CN114248905B true CN114248905B (en) | 2024-04-26 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111392164.XA Active CN114248905B (en) | 2021-11-23 | 2021-11-23 | Composite wing spar and forming method thereof |
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| CN (1) | CN114248905B (en) |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN206265287U (en) * | 2016-12-23 | 2017-06-20 | 厦门欧势复材科技有限公司 | A kind of unmanned plane carbon fibre composite spar |
| CN109484624A (en) * | 2018-11-29 | 2019-03-19 | 中国商用飞机有限责任公司北京民用飞机技术研究中心 | A kind of technique for aircraft composite wing spar and wing root area connection structure |
| CN111278736A (en) * | 2017-10-27 | 2020-06-12 | 萨维尔股份公司 | Method for assembling portions of an aircraft wing |
| CN214566114U (en) * | 2021-03-12 | 2021-11-02 | 中航西飞民用飞机有限责任公司 | Whole wing spar structure of aircraft |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20200086970A1 (en) * | 2018-09-18 | 2020-03-19 | The Boeing Company | Composite fabric wing spar with interleaved tape cap plies |
-
2021
- 2021-11-23 CN CN202111392164.XA patent/CN114248905B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN206265287U (en) * | 2016-12-23 | 2017-06-20 | 厦门欧势复材科技有限公司 | A kind of unmanned plane carbon fibre composite spar |
| CN111278736A (en) * | 2017-10-27 | 2020-06-12 | 萨维尔股份公司 | Method for assembling portions of an aircraft wing |
| CN109484624A (en) * | 2018-11-29 | 2019-03-19 | 中国商用飞机有限责任公司北京民用飞机技术研究中心 | A kind of technique for aircraft composite wing spar and wing root area connection structure |
| CN214566114U (en) * | 2021-03-12 | 2021-11-02 | 中航西飞民用飞机有限责任公司 | Whole wing spar structure of aircraft |
Also Published As
| Publication number | Publication date |
|---|---|
| CN114248905A (en) | 2022-03-29 |
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