CN109228394B - Rapid forming method for composite material fuselage reinforcing rib - Google Patents
Rapid forming method for composite material fuselage reinforcing rib Download PDFInfo
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- CN109228394B CN109228394B CN201810978281.6A CN201810978281A CN109228394B CN 109228394 B CN109228394 B CN 109228394B CN 201810978281 A CN201810978281 A CN 201810978281A CN 109228394 B CN109228394 B CN 109228394B
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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
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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
- B29C70/54—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
- 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
- B29C70/543—Fixing the position or configuration of fibrous reinforcements before or during moulding
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- 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
- B29L2031/3082—Fuselages
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Composite Materials (AREA)
- Mechanical Engineering (AREA)
- Textile Engineering (AREA)
- Moulding By Coating Moulds (AREA)
- Road Paving Structures (AREA)
Abstract
The invention belongs to the technical field of resin-based composite material forming, and relates to a method for quickly forming a composite material fuselage reinforcing rib. In the paving and pasting process of the composite material reinforcing rib, the web plates of the 0-degree layers are automatically paved on the template, then the 0-degree preformed body of the automatic paving is transferred to the forming tool, the paving and pasting efficiency is improved, the paving and pasting period of the 0-degree layers is effectively shortened, the splicing seam of the next 0-degree layer is staggered with the splicing seam of the previous 0-degree layer, and the surface quality of the composite material fuselage reinforcing rib is ensured; in the forming process of the composite material machine body reinforcing rib, the soft cover plate is omitted, and the thickness and the surface quality of the composite material machine body reinforcing rib are guaranteed. The forming method of the composite material fuselage reinforcing rib provided by the invention ensures the forming quality of the composite material reinforcing disc, simplifies the forming process, shortens the manufacturing period and reduces the manufacturing cost.
Description
Technical Field
The invention belongs to the technical field of resin-based composite material forming, and relates to a method for quickly forming a composite material fuselage reinforcing rib.
Background
The composite material is widely used in the field of aviation due to a series of characteristics of high specific strength, high specific modulus, designability, fatigue resistance and the like. From the non-bearing member to the secondary bearing member and from the secondary bearing member to the main bearing member. The composite material reinforcing rib (comprising a partition frame and a corner piece) of the fuselage is a main bearing component, and is characterized in that the 0-degree layering direction of the component is the circumferential direction and is not the conventional fixed direction, and the characteristic provides new challenges for the thickness uniformity, the surface quality and the appearance tolerance of the composite material reinforcing rib. The conventional forming method of the composite material fuselage reinforcing rib comprises the following steps: manufacturing a reinforcing rib forming tool, paving and solidifying a reinforcing rib process dummy part, modifying the shape of the process dummy part, paving and solidifying a soft cover plate and paving and solidifying a reinforcing rib formal part. The thickness and the surface quality of the composite material reinforcing rib manufactured by the conventional forming method hardly meet the design requirements, and the method has the advantages of complex process, long manufacturing period and high manufacturing cost.
The invention aims to provide a method for molding a composite material fuselage reinforcing rib, which ensures the molding quality (especially the thickness and the surface quality of parts) of the composite material reinforcing rib, simplifies the molding process, shortens the manufacturing period and reduces the manufacturing cost.
The technical solution of the invention is as follows:
(1) designing and manufacturing a reinforcing rib forming tool according to the manufacturing technical index of the composite material fuselage reinforcing rib;
(2) preparing an automatic prepreg blanking file and a laser projection file according to the manufacturing technical indexes of the composite material fuselage reinforcing rib, adopting non-0-degree laying layers of the edge strip and the web plate of the composite material fuselage reinforcing rib and 0-degree laying layers of the edge strip for automatic blanking of prepreg, and manually laying on a forming tool under the assistance of laser projection;
(3) the web plates of the 0-degree layer are automatically laid on a template according to an automatic wire laying file by using prepreg of an automatic wire laying process, and the 0-degree layer pre-forming body formed by automatic wire laying is transferred to a forming tool, wherein the splicing seam of the next 0-degree layer must be staggered with the splicing seam of the previous 0-degree layer;
(4) repeating the steps 2 and 3, and vacuumizing once every 1-3 layers of paving until all the layers are paved; then packaging, curing, subsequent processing and detecting are carried out.
The section of the composite material fuselage reinforcing rib is Z-shaped, L-shaped or J-shaped.
The staggered distance between the splicing seam of the next 0-degree laying layer and the splicing seam of the previous 0-degree laying layer is half of the width of the prepreg of the automatic fiber laying process.
The invention has the advantages and beneficial effects
The invention greatly simplifies the forming method of the composite material fuselage reinforcing rib, effectively ensures the forming quality of the composite material reinforcing rib, simplifies the forming process, simplifies the processes of manufacturing the reinforcing rib forming tool, paving and solidifying the reinforcing rib process dummy piece, shaping the process dummy piece, paving and solidifying the soft cover plate, paving and solidifying the reinforcing rib formal piece and the like in the conventional forming method into the manufacturing of the reinforcing rib forming tool and the paving and solidifying of the reinforcing rib formal piece, and greatly reduces the whole manufacturing period and workload; in the paving and pasting process of the composite material reinforcing rib, the web plates of the 0-degree layers are automatically paved on the template, then the 0-degree preformed body of the automatic paving is transferred to a forming tool, the paving and pasting efficiency is improved, the paving and pasting period of the 0-degree layers is effectively shortened, the splicing seam of the next 0-degree layer is staggered with the splicing seam of the previous 0-degree layer, and the surface quality of the composite material fuselage reinforcing rib is ensured; in the forming process of the composite material machine body reinforcing rib, the soft cover plate is omitted, and the thickness and the surface quality of the composite material machine body reinforcing rib are guaranteed. The forming method of the composite material fuselage reinforcing rib provided by the invention ensures the forming quality of the composite material reinforcing disc, simplifies the forming process, shortens the manufacturing period and reduces the manufacturing cost.
Drawings
FIG. 1 is a schematic illustration of a method of forming a conventional composite fuselage stiffener;
FIG. 2 is a schematic view of a method of forming a composite fuselage stiffener according to the present disclosure;
FIG. 3 is a schematic structural view of a Z-shape of a composite fuselage stiffener; wherein FIG. 3a is a schematic view of the zero degree orientation and FIG. 3b is a schematic view of the web and bead positions;
FIG. 4 is a schematic view of a composite fuselage stiffener L-shape; wherein FIG. 4a is a schematic view of the zero degree orientation and FIG. 4b is a schematic view of the web and bead positions;
fig. 5 is a structural schematic diagram of a forming tool for the composite material fuselage reinforcing rib.
Detailed Description
(1) Designing (selecting a material of a tool, selecting a structural form of the tool, determining a manufacturing tolerance and the like) and manufacturing a reinforcing rib forming tool according to manufacturing technical indexes (appearance requirement, internal quality, assembly requirement and the like) of the composite material fuselage reinforcing rib;
(2) the non-0-degree ply and 0-degree ply flanges generate an automatic prepreg blanking file and a laser projection file by using composite ply software (such as Composites Design module and FiberSIM of CATIA) according to the manufacturing technical requirements (ply angle tolerance, butt seam processing mode and repeated arrangement of the same ply angle) of the composite material fuselage reinforcing rib; preparing an automatic fiber-laying file for a web plate with 0-degree layers according to the manufacturing technical requirements of the composite material fuselage reinforcing rib, wherein the splicing seam of the next 0-degree layer must be staggered with the splicing seam of the previous 0-degree layer, and the splicing seams must not coincide with each other, and the staggered automatic fiber-laying technology is suggested to be half of the width of a prepreg for an automatic fiber-laying process;
(3) importing the prepreg automatic blanking file into a prepreg automatic blanking machine (such as a GEBER cutter and an Esubject blanking machine), and importing the laser projection file into a laser projector (such as a Virtek laser projector and a LAP projector); carrying out prepreg blanking on the edge strips of the non-0-degree layer and the 0-degree layer according to an automatic blanking file, and manually paving and pasting the prepreg on a reinforcing rib forming tool according to laser projection; and (3) introducing the automatic fiber laying file into an automatic fiber laying machine, automatically laying a web preformed body with 0-degree laying layers on a template, and then transferring the web preformed body with 0-degree laying layers of the automatic fiber laying to a forming tool.
(4) Repeating the steps 2 and 3, and vacuumizing once every 1-3 layers of paving until all the layers are paved;
(5) packaging and curing according to the process specification;
(6) and (4) demolding the solidified composite material machine body reinforcing rib, machining, detecting internal quality, detecting thickness, detecting surface quality and the like.
Examples
The cross-section of certain combined material fuselage strengthening rib is the Z shape, comprises cap 2 and web 3, and length is about 5000mm, and the width is about 150mm, and highly is about 100mm, and thickness is about 2.24mm, is formed by 12 layers of unidirectional tape prepreg paving, and it spreads the ply angle and is: [+45/+45/0/0/-45/90]SAnd the thickness of the Z-shaped composite material reinforcing rib is required to be uniform and the surface is required to be smooth along the circumferential direction along the 0-degree direction 1.
The implementation of the invention is as follows:
(1) according to the manufacturing technical indexes of the composite material fuselage reinforcing rib: the requirements on the degree of sticking, the surface quality and the thickness of parts are high, factors such as the uniformity of a temperature field, the manufacturability of paving and sticking, the overall dimension and the like are considered, the reinforcing rib forming tool 4 is made of an invar material, a frame type structure is adopted, and two occupied spaces 5 are designed on each forming tool.
(2) According to the manufacturing technical requirements of the composite material fuselage reinforcing rib (the ply angle tolerance is +/-3 degrees, the width direction butt joint of the unidirectional prepreg and the 0-degree ply direction definition and the like), a flange 2 of the non-0-degree ply and 0-degree ply of the composite material fuselage reinforcing rib is used for generating an automatic blanking file and a laser projection file of the prepreg by using composite material ply software FiberSIM, the blanking width of the prepreg is determined to be 140mm according to the manufacturability analysis and the convenience of actual manufacturing, the blanking width of the 0-degree ply prepreg is added with allowance blanking according to the width of the flange 2, the automatic blanking file and the laser projection file are generated by using the composite material ply software FiberSIM, an automatic wirelaying file of a 0-degree ply web 3 is prepared by using a CADFiber according to the manufacturing technical requirements of the composite material fuselage reinforcing rib, wherein the splicing seam of the next 0-degree ply is staggered with the splicing seam of the previous 0-degree ply, the overlapping is not required;
(3) introducing the prepreg automatic blanking file generated in the step 2 into a Gerber cutter prepreg automatic blanking machine (Gerber Technology, USA), introducing a laser projection file into a laser projector Virtek LaserEdge, performing prepreg blanking on a non-0-degree paving layer (+ 45-degree, 45-degree and 90-degree paving layers) and a 0-degree paving layer edge strip 2 according to the automatic blanking file, and performing manual prepreg paving on a reinforcing rib forming tool according to laser projection; importing the automatic fiber laying file into an automatic fiber laying machine, automatically laying a web 3 preformed body with 0-degree laying layers on a template, and transferring the preformed body with automatic fiber laying to a forming tool;
(4) repeating the steps 2 and 3, and vacuumizing once every 1-3 layers of paving until all the layers are paved and transferred;
(5) auxiliary materials such as a non-porous isolating membrane, an air-permeable felt and a vacuum bag are respectively placed on the vacuumized reinforcing ribs of the fuselage, and the auxiliary materials such as the non-porous isolating membrane, the air-permeable felt and the vacuum bag cannot be bridged;
(6) and (4) demolding the solidified composite material machine body reinforcing rib, machining, detecting internal quality, detecting thickness, detecting surface quality and the like.
Claims (3)
1. A method for quickly forming a reinforcing rib of a composite material machine body is characterized in that,
(1) designing and manufacturing a reinforcing rib forming tool according to the manufacturing technical index of the composite material fuselage reinforcing rib;
(2) preparing an automatic prepreg blanking file and a laser projection file according to the manufacturing technical indexes of the composite material fuselage reinforcing rib, adopting non-0-degree laying layers of the edge strip and the web plate of the composite material fuselage reinforcing rib and 0-degree laying layers of the edge strip for automatic blanking of prepreg, and manually laying on a forming tool under the assistance of laser projection;
(3) the web plates of the 0-degree layer are automatically laid on a template according to an automatic wire laying file by using prepreg of an automatic wire laying process, and the 0-degree layer pre-forming body formed by automatic wire laying is transferred to a forming tool, wherein the splicing seam of the next 0-degree layer must be staggered with the splicing seam of the previous 0-degree layer;
(4) repeating the steps 2 and 3, and vacuumizing once every 1-3 layers of paving until all the layers are paved; then packaging, curing, subsequent processing and detecting are carried out;
wherein, the soft cover plate is not used in the molding process, and the prepreg using the automatic wire laying process is used for 0-degree laying of the reinforcing ribs.
2. The method of claim 1 wherein the composite fuselage stiffener has a Z, L or J-shaped cross-section.
3. The method for forming the composite material fuselage reinforcing rib as claimed in claim 1, wherein the distance between the splicing seam of the next 0-degree ply and the splicing seam of the previous 0-degree ply is half of the width of the prepreg of the automatic filament laying process.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810978281.6A CN109228394B (en) | 2018-08-24 | 2018-08-24 | Rapid forming method for composite material fuselage reinforcing rib |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810978281.6A CN109228394B (en) | 2018-08-24 | 2018-08-24 | Rapid forming method for composite material fuselage reinforcing rib |
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| Publication Number | Publication Date |
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| CN109228394A CN109228394A (en) | 2019-01-18 |
| CN109228394B true CN109228394B (en) | 2020-12-29 |
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| CN201810978281.6A Active CN109228394B (en) | 2018-08-24 | 2018-08-24 | Rapid forming method for composite material fuselage reinforcing rib |
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Families Citing this family (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110978561A (en) * | 2019-12-30 | 2020-04-10 | 中国航空制造技术研究院 | Preparation method of fiber preform of composite material cover body with full-layering structure |
| CN111361179A (en) * | 2020-03-30 | 2020-07-03 | 西安交通大学 | Thermoplastic composite material forming process suitable for complex large curvature |
| CN112810182B (en) * | 2020-12-29 | 2021-11-30 | 江苏新扬新材料股份有限公司 | Forming method of composite material cylindrical support |
| CN114646923A (en) * | 2022-01-04 | 2022-06-21 | 威海光威复合材料股份有限公司 | Radar part that inside and outside structure switches on |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104484516A (en) * | 2014-12-04 | 2015-04-01 | 江苏恒神纤维材料有限公司 | Method of laying prepreg by aid of trajectory planning software |
| CN105172161A (en) * | 2015-08-21 | 2015-12-23 | 航天材料及工艺研究所 | Automatic fiber placement forming method for grid skin structure with part concave structure |
| CN106239941A (en) * | 2016-08-29 | 2016-12-21 | 中国航空工业集团公司基础技术研究院 | The projection of a kind of material prepreg and laying method |
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- 2018-08-24 CN CN201810978281.6A patent/CN109228394B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104484516A (en) * | 2014-12-04 | 2015-04-01 | 江苏恒神纤维材料有限公司 | Method of laying prepreg by aid of trajectory planning software |
| CN105172161A (en) * | 2015-08-21 | 2015-12-23 | 航天材料及工艺研究所 | Automatic fiber placement forming method for grid skin structure with part concave structure |
| CN106239941A (en) * | 2016-08-29 | 2016-12-21 | 中国航空工业集团公司基础技术研究院 | The projection of a kind of material prepreg and laying method |
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