CN113635570A - Preparation method of unmanned aerial vehicle composite material body structure - Google Patents
Preparation method of unmanned aerial vehicle composite material body structure Download PDFInfo
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- CN113635570A CN113635570A CN202110971672.7A CN202110971672A CN113635570A CN 113635570 A CN113635570 A CN 113635570A CN 202110971672 A CN202110971672 A CN 202110971672A CN 113635570 A CN113635570 A CN 113635570A
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- prepreg
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- aerial vehicle
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- 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
-
- 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
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C2001/0054—Fuselage structures substantially made from particular materials
- B64C2001/0072—Fuselage structures substantially made from particular materials from composite materials
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- 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
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Chemical & Material Sciences (AREA)
- Composite Materials (AREA)
- Aviation & Aerospace Engineering (AREA)
- Laminated Bodies (AREA)
- Moulding By Coating Moulds (AREA)
Abstract
The invention discloses a preparation method of a composite material body structure of an unmanned aerial vehicle, wherein the composite material body structure comprises a body beam, a body skin, a wing rib, a wing skin, a tail wing rib, a head skin and a tail wing skin, and the preparation steps of the composite material body structure are as follows: coating a degumming agent on the surface of the steel mould; laying a prepreg; carrying out curing molding; trimming and trimming. The invention uses the composite material as the unmanned aerial vehicle body structure, thereby reducing the structural weight of the unmanned aerial vehicle and improving the flight speed, the cruising ability and the maneuverability of the unmanned aerial vehicle.
Description
Technical Field
The invention relates to the field of composite materials, in particular to a preparation method of an unmanned aerial vehicle composite material body structure.
Background
Unmanned aerial vehicles are unmanned planes, and along with the rapid development of electronic technologies and materials, unmanned aerial vehicles are widely applied to the fields of communication relay, aerial photography, resource exploration, military and the like. Unmanned aerial vehicle mainly comprises fuselage, wing, fin, structural component such as.
Unmanned aerial vehicle organism structure is unmanned aerial vehicle's truck and atress basis, not only will fix and support unmanned aerial vehicle's other parts, connects into a whole with whole unmanned aerial vehicle, still bears the load that each adapting unit transmitted, bears load at the inside equipment of fuselage, task load and gravity and inertia itself.
Most of structures of the existing unmanned aerial vehicle are made of alloy, and the flight speed and the maneuverability of the unmanned aerial vehicle are seriously influenced due to the weight of the alloy.
Disclosure of Invention
Technical problem to be solved by the invention
The invention aims to provide a method for preparing a composite material body structure of an unmanned aerial vehicle, which aims to solve the problems in the background art that most of existing unmanned aerial vehicles are made of alloy, and the flight speed and the maneuverability of the unmanned aerial vehicle are seriously influenced due to the weight of the alloy.
Technical scheme
In order to achieve the purpose, the invention provides the following technical scheme: a preparation method of an unmanned aerial vehicle composite material body structure comprises a body beam, a body skin, a wing rib, a wing skin, a tail rib, a machine head skin and a tail wing skin, and the preparation steps of the composite material body structure are as follows:
coating a degumming agent on the surface of the steel mould;
laying a prepreg;
carrying out curing molding;
trimming and trimming.
Preferably, the manufacturing steps of the fuselage beam, the wing ribs and the tail wing rib are as follows: smearing a degumming agent on the surface of a rigid mould, starting a laying operation after the degumming agent is volatilized, laying a prepreg on a steel mould, closing the mould, putting the mould into an oven for curing, opening the steel mould after curing, and trimming and shaping.
The manufacturing steps of the fuselage skin, the wing skin, the nose skin and the empennage skin are as follows: smearing a degumming agent on the surface of a rigid mould, volatilizing the degumming agent, carrying out layer laying operation, laying the prepreg on a steel mould, laying a porous isolation material, a glue absorbing material, a pressure equalizing plate, a breathable material and a vacuum bag, vacuumizing, putting into an oven for curing, opening the steel mould after curing, and trimming and shaping.
Preferably, the fuselage beam, the fuselage skin, the wing ribs, the wing skin and the tail wing ribs are of carbon fiber body structures, and the nose skin and the tail wing skin are of glass fiber body structures.
Preferably, the prepreg of the carbon fiber body structure is formed by an oven curing process, and the parameters of the oven curing process are as follows: heating to 120-130 deg.C at a speed of 1-3 deg.C/min, maintaining at 0.08-0.12MPa for 1-2h, and naturally cooling to below 60 deg.C.
Preferably, the prepreg of the glass fiber body structure is formed by an oven curing process, and the oven curing process parameters are as follows: heating to 120-130 deg.C at a speed of 1-3 deg.C/min, maintaining at 0.08-0.12MPa for 1-2h, and naturally cooling to below 60 deg.C.
Preferably, when the carbon fiber prepreg and the glass fiber prepreg are co-cured, the curing process parameters of the glass fiber prepreg are used as the standard.
Preferably, the area density of the prepreg of the carbon fiber prepreg is 350-390g/m2Twill fabric with 37-42% resin content, 0.2-0.25mm single layer thickness, 120-140 deg.c gel temperature and tensile strength550-650MPa, 500-600MPa and 50-60 MPa.
Preferably, the prepreg areal density of the glass fiber prepreg is 380-420g/m2The plain weave fabric has the resin content of 37-42 percent, the single-layer thickness of 0.2-0.25mm, the tensile strength of 450-550MPa, the bending strength of 450-500MPa and the interlaminar shear strength of 50-55 MPa.
Preferably, after the prepreg is cured and molded, tensile and compressive moduli in the directions of 0 DEG and 90 DEG of the fabric laminate are not less than 60GPa, and tensile and compressive strengths are not less than 650 MPa.
Preferably, a paving and pasting process is adopted to pave the prepreg, the paving starting point is selected, the paving coordinate origin in the model is used as the basis, the paving tolerance requirement is met, during paving, the paving must be continuous along the fiber direction, the prepreg cannot be randomly cut off and lapped, when the width of the prepreg is insufficient, splicing is allowed, the splicing seams are staggered, three sides of a part meet, or the shearing opening is allowed in a sharply-turning area, and the shearing openings are staggered.
Advantageous effects
Compared with the prior art, the invention has the beneficial effects that: the invention uses the composite material as the unmanned aerial vehicle body structure, thereby reducing the structural weight of the unmanned aerial vehicle and improving the flight speed, the cruising ability and the maneuverability of the unmanned aerial vehicle.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments of the invention without making creative efforts, shall fall within the protection scope of the invention.
Example 1
The invention provides a technical scheme that: the utility model provides an unmanned aerial vehicle combined material organism structure preparation method, includes combined material organism structure, combined material organism structure includes carbon-fibre composite material organism structure and glass fiber composite material organism structure, carbon-fibre composite material organism structure includes fuselage roof beam, fuselage skin, wing rib, wing skin and tail wing rib, glass fiber composite material organism structure includes aircraft nose skin and fin skin, combined material organism structure adopts prepreg spread layer structure, through the preparation of solidification process, unmanned aerial vehicle has the effect that alleviates unmanned aerial vehicle structure weight through using combined material organism structure, and then improves unmanned aerial vehicle flying speed, duration and mobility.
The manufacturing steps of the fuselage beam, the wing ribs and the tail wing rib are as follows: firstly, coating a degumming agent on the surface of a steel mould, starting a layer laying operation after the degumming agent is volatilized, laying a prepreg on a steel mould, then closing the mould, then putting the mould into an oven for curing, opening the steel mould after the curing is finished, and trimming and repairing the mould, wherein carbon fiber is mainly special fiber consisting of carbon elements, the carbon content of the carbon fiber is different according to different types and is generally more than 90%. The carbon fiber has the characteristics of a general carbon material, such as high temperature resistance, friction resistance, electric conduction, heat conduction, corrosion resistance and the like, but has remarkable anisotropy in appearance, is soft, can be processed into various fabrics and shows high strength along the fiber axis direction, unlike the general carbon material.
The manufacturing steps of the fuselage skin, the wing skin, the nose skin and the empennage skin are as follows: the method comprises the steps of firstly coating a degumming agent on the surface of a rigid mold, starting a layering operation after the degumming agent is volatilized, laying a prepreg on a steel mold, then laying a porous isolation material, a glue absorbing material, a pressure equalizing plate, a breathable material and a vacuum bag in sequence, vacuumizing, putting the steel mold into an oven for curing after vacuumizing, opening the steel mold after curing, and trimming and repairing the steel mold.
The prepreg of the carbon fiber body structure is formed by adopting an oven curing and forming process, and the parameters of the oven curing and forming process are as follows: heating to 120-130 deg.C at 1-3 deg.C/min, maintaining at 0.08-0.12MPa for 1-2h, and naturally cooling to below 60 deg.C.
The prepreg of the glass fiber body structure is formed by adopting an oven curing process, and the oven curing process parameters are as follows: heating to 120-130 deg.C at 1-3 deg.C/min, maintaining at 0.08-0.12MPa for 1-2h, and naturally cooling to below 60 deg.C.
When the carbon fiber prepreg and the glass fiber prepreg are cured together, the curing process parameters of the glass fiber prepreg are used as the standard.
The surface density of the carbon fiber prepreg is 350-390g/m2The twill fabric has the resin content of 37-42 percent, the single-layer thickness of 0.2-0.25mm, the gel temperature of 120-.
The surface density of the prepreg of the glass fiber prepreg is 380-420g/m2The plain weave fabric has the resin content of 37-42 percent, the single-layer thickness of 0.2-0.25mm, the tensile strength of 450-550MPa, the bending strength of 450-500MPa and the interlaminar shear strength of 50-55 MPa.
Laying the prepreg by adopting a laying and pasting process, wherein the laying tolerance is not more than +/-5 degrees; the selection of the laying starting point meets the laying tolerance requirement by taking the laying coordinate origin in the model as the basis; in the part model, a model entity is the only basis of the boundary of the composite material part; the boundary of the composite material layering is only used for reference, and the boundary is allowed to be adjusted by combining the requirement of the process cutting allowance; when in paving and pasting, the fiber must be continuous along the fiber direction, and can not be cut off and lapped randomly; when the width of the prepreg is not enough, splicing is allowed, and a gap larger than 1mm or an overlap larger than 1mm in width is not needed at the splicing position; the abutted seams should be staggered at 4-layer angles, and the mutual staggered distance is more than or equal to 20 mm. The three sides of the part are converged or the part is sharply turned, the cut is allowed to be staggered, the cut is staggered with 4 layers of angles, the staggered distance is larger than or equal to 10mm, the adjacent 4 layers of angles are staggered with the staggered distance larger than or equal to 10mm, the tip part of the cut is required to be rounded, and the radius of the fillet is 1-5 mm.
After the prepreg is cured and molded, the tensile and compressive moduli in the directions of 0 DEG and 90 DEG of the fabric laminated plate are not less than 60GPa, and the tensile and compressive strengths are not less than 650 MPa.
Example 2
In this embodiment, including composite material organism structure, composite material organism structure includes carbon-fibre composite material organism structure and glass fiber composite material organism structure, carbon-fibre composite material organism structure includes fuselage roof beam, fuselage skin, wing rib, wing skin and tail wing rib, glass fiber composite material organism structure includes aircraft nose skin and fin skin, composite material organism structure adopts prepreg to spread the layer structure, forms through curing moulding technology preparation.
The manufacturing steps of the fuselage beam, the wing ribs and the tail wing rib are as follows: firstly, coating a degumming agent on the surface of a steel mould, starting the laying operation after the degumming agent is volatilized, laying a prepreg on a steel mould, then closing the mould, then putting the mould into an autoclave for curing, opening the steel mould after the curing is finished, and trimming and shaping.
The manufacturing steps of the fuselage skin, the wing skin, the nose skin and the empennage skin are as follows: firstly, coating a degumming agent on the surface of a steel mould, starting a layer laying operation after the degumming agent is volatilized, laying a prepreg on a steel mould, then laying a porous isolation material, a glue absorbing material, a pressure equalizing plate, a breathable material and a vacuum bag in sequence, vacuumizing, putting the steel mould into an autoclave for curing, opening the steel mould after curing, and trimming and repairing the mould.
The prepreg of the carbon fiber body structure is formed by adopting an autoclave curing process, and autoclave curing process parameters are as follows: heating to 120-130 deg.C at 1-3 deg.C/min, pressurizing to 0.6MPa, maintaining for 1-2 hr, and naturally cooling to below 60 deg.C; the mould is a steel mould.
The prepreg of the glass fiber body structure adopts an autoclave curing molding process, and autoclave curing process parameters are as follows: heating to 120-130 deg.C at 1-3 deg.C/min, pressurizing to 0.25-0.35MPa, holding for 1-2 hr, and naturally cooling to below 60 deg.C.
When the carbon fiber prepreg and the glass fiber prepreg are cured together, the curing process parameters of the glass fiber prepreg are used as the standard.
The surface density of the carbon fiber prepreg is 350-390g/m2The twill fabric has the resin content of 37-42 percent, the single-layer thickness of 0.2-0.25mm, the gel temperature of 120-The strength is 50-60 MPa.
The surface density of the prepreg of the glass fiber prepreg is 380-420g/m2The plain weave fabric has the resin content of 37-42 percent, the single-layer thickness of 0.2-0.25mm, the tensile strength of 450-550MPa, the bending strength of 450-500MPa and the interlaminar shear strength of 50-55 MPa.
Laying the prepreg by adopting a laying and pasting process, wherein the laying tolerance is not more than +/-5 degrees; the selection of the laying starting point meets the laying tolerance requirement by taking the laying coordinate origin in the model as the basis; in the part model, a model entity is the only basis of the boundary of the composite material part; the boundary of the composite material layering is only used for reference, and the boundary is allowed to be adjusted by combining the requirement of the process cutting allowance; when in paving and pasting, the fiber must be continuous along the fiber direction, and can not be cut off and lapped randomly; when the width of the prepreg is not enough, splicing is allowed, and a gap larger than 1mm or an overlap larger than 1mm in width is not needed at the splicing position; the abutted seams should be staggered at 4-layer angles, and the mutual staggered distance is more than or equal to 20 mm. The three sides of the part are converged or the part is sharply turned, the cut is allowed to be staggered, the cut is staggered with 4 layers of angles, the staggered distance is larger than or equal to 10mm, the adjacent 4 layers of angles are staggered with the staggered distance larger than or equal to 10mm, the tip part of the cut is required to be rounded, and the radius of the fillet is 1-5 mm.
After the prepreg is cured and molded, the tensile and compressive moduli in the directions of 0 DEG and 90 DEG of the fabric laminated plate are not less than 60GPa, and the tensile and compressive strengths are not less than 650 MPa.
If the traditional unmanned aerial vehicle is an all-metal part, the body structure of the traditional unmanned aerial vehicle is 60kg, and is replaced by partial composite materials, the weight is reduced to 40kg, the flying speed of the traditional unmanned aerial vehicle is about 400 km/h, and the speed of the unmanned aerial vehicle manufactured by the invention can reach subsonic speed-850-900 km/h.
The above description is only for the specific embodiments of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art can easily conceive of the changes or substitutions within the technical scope of the present invention.
Claims (10)
1. The preparation method of the unmanned aerial vehicle composite material body structure is characterized by comprising the following steps: the composite material body structure comprises a body beam, a body skin, a wing rib, a wing skin, a tail rib, a nose skin and a tail wing skin, and the manufacturing steps of the composite material body structure are as follows:
coating a degumming agent on the surface of the steel mould;
laying a prepreg;
carrying out curing molding;
trimming and trimming.
2. The method for preparing the airframe structure of unmanned aerial vehicle as defined in claim 1, wherein: the manufacturing steps of the fuselage beam, the wing ribs and the tail wing rib are as follows: smearing a degumming agent on the surface of a rigid mould, laying a prepreg on a steel mould after the degumming agent is volatilized, closing the mould, putting the mould into an oven for curing, opening the steel mould after curing, and trimming and shaping;
the manufacturing steps of the fuselage skin, the wing skin, the nose skin and the empennage skin are as follows: smearing a degumming agent on the surface of a rigid mould, carrying out layer laying operation after the degumming agent is volatilized, laying a prepreg on a steel mould, laying a porous isolation material, a glue absorbing material, a pressure equalizing plate, a breathable material and a vacuum bag, vacuumizing, putting the steel mould into an oven for curing, opening the steel mould after curing, and trimming and shaping.
3. The method for preparing the airframe structure of unmanned aerial vehicle as defined in claim 1, wherein: the aircraft body beam, the aircraft body skin, the wing ribs, the wing skin and the tail wing ribs are of carbon fiber body structures, and the aircraft head skin and the tail wing skin are of glass fiber body structures.
4. The method for preparing the airframe structure of unmanned aerial vehicle as claimed in claim 2, wherein: the prepreg in the carbon fiber body structure manufacturing step adopts an oven curing molding process, and the parameters of the oven curing molding process are as follows: heating to 120-130 deg.C at a speed of 1-3 deg.C/min, maintaining at 0.08-0.12MPa for 1-2h, and naturally cooling to below 60 deg.C.
5. The method for preparing the airframe structure of unmanned aerial vehicle as claimed in claim 2, wherein: the prepreg in the step of manufacturing the glass fiber body structure adopts an oven curing molding process, and the parameters of the oven curing process are as follows: heating to 120-130 deg.C at a speed of 1-3 deg.C/min, maintaining at 0.08-0.12MPa for 1-2h, and naturally cooling to below 60 deg.C.
6. The method for preparing the airframe structure of unmanned aerial vehicle as defined in claim 1, wherein: when the carbon fiber prepreg and the glass fiber prepreg are cured together, the curing process parameters of the glass fiber prepreg are used as the standard.
7. The method for preparing the airframe structure of unmanned aerial vehicle as claimed in claim 5, wherein: the surface density of the carbon fiber prepreg is 350-390g/m2The twill fabric has the resin content of 37-42 percent, the single-layer thickness of 0.2-0.25mm, the gel temperature of 120-.
8. The method for preparing the airframe structure of unmanned aerial vehicle as claimed in claim 5, wherein: the surface density of the prepreg of the glass fiber prepreg is 380-420g/m2The plain weave fabric has the resin content of 37-42 percent, the single-layer thickness of 0.2-0.25mm, the tensile strength of 450-550MPa, the bending strength of 450-500MPa and the interlaminar shear strength of 50-55 MPa.
9. The method for preparing the airframe structure of unmanned aerial vehicle as defined in claim 1, wherein: after the prepreg is cured and molded, the prepreg is stretched and compressed along the directions of 0 degrees and 90 degrees of the fabric laminated plate, the modulus is not less than 60GPa, and the tensile and compressive strength is not lower than 650 MPa.
10. The method for preparing the airframe structure of unmanned aerial vehicle as defined in claim 1, wherein: the prepreg is paved by adopting a paving and pasting process, a paving starting point is selected, the paving coordinate origin in the model is taken as the basis, the paving tolerance requirement is met, the prepreg must be continuous along the fiber direction during paving and pasting, the prepreg cannot be randomly cut off and lapped, when the width of the prepreg is insufficient, splicing is allowed, the splicing seams are staggered, three sides of the part meet, or the area with sharp turning is allowed to be cut, and the cutting seams are staggered mutually.
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| CN202110971672.7A CN113635570A (en) | 2021-08-24 | 2021-08-24 | Preparation method of unmanned aerial vehicle composite material body structure |
| CN202210372688.0A CN114919205A (en) | 2021-08-24 | 2022-04-11 | Preparation method of unmanned aerial vehicle composite material body structure |
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| CN102795338B (en) * | 2012-07-27 | 2014-10-08 | 北京卫星制造厂 | Micro unmanned aerial vehicle carbon fiber rotor wing and preparation method thereof |
| CN105416567A (en) * | 2015-11-13 | 2016-03-23 | 中国人民解放军国防科学技术大学 | Skin, unmanned aerial vehicle wing, manufacturing method of unmanned aerial vehicle wing, empennage and manufacturing method of empennage |
| CN107215039B (en) * | 2017-06-07 | 2023-03-14 | 国电联合动力技术有限公司 | Sandwich composite material and preparation method thereof |
| US10391722B1 (en) * | 2017-07-12 | 2019-08-27 | University Of Iceland | Method of producing aerofoils |
| CN107628232A (en) * | 2017-08-11 | 2018-01-26 | 精功（绍兴）复合材料技术研发有限公司 | A kind of composite unmanned airplane empennage and its manufacture method |
| US10640197B1 (en) * | 2017-10-31 | 2020-05-05 | Wisk Aero Llc | Composite airfoil with rolled fiberglass reinforced leading edge |
| CN108262993A (en) * | 2018-02-02 | 2018-07-10 | 上海晋飞碳纤科技股份有限公司 | A kind of ply angles and integral forming process of civil aircraft rudder |
| CN108959744B (en) * | 2018-06-21 | 2022-11-15 | 西安爱生技术集团公司 | Manufacturing method of composite wing box of small and medium-sized unmanned aerial vehicle |
| CN110193956B (en) * | 2019-07-03 | 2021-05-18 | 西安爱生技术集团公司 | Small unmanned aerial vehicle composite material vertical tail wing forming process method |
| CN110193955B (en) * | 2019-07-03 | 2021-05-18 | 西安爱生技术集团公司 | Process forming method for composite outer wing of small and medium-sized unmanned aerial vehicle |
| CN111113949A (en) * | 2019-12-26 | 2020-05-08 | 航天神舟飞行器有限公司 | Forming die and method for double lug inserting pieces of power cabin of solar unmanned aerial vehicle |
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