CN117162547A - Forming method for continuous layering of composite material blade of aero-engine - Google Patents
Forming method for continuous layering of composite material blade of aero-engine Download PDFInfo
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- CN117162547A CN117162547A CN202310983831.4A CN202310983831A CN117162547A CN 117162547 A CN117162547 A CN 117162547A CN 202310983831 A CN202310983831 A CN 202310983831A CN 117162547 A CN117162547 A CN 117162547A
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- layering
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- 239000002131 composite material Substances 0.000 title claims abstract description 42
- 238000000034 method Methods 0.000 title claims abstract description 40
- 238000000465 moulding Methods 0.000 claims abstract description 29
- 239000012779 reinforcing material Substances 0.000 claims description 32
- 239000000835 fiber Substances 0.000 claims description 19
- 239000000463 material Substances 0.000 claims description 19
- 229920005989 resin Polymers 0.000 claims description 9
- 239000011347 resin Substances 0.000 claims description 9
- 230000002787 reinforcement Effects 0.000 claims description 7
- 239000003822 epoxy resin Substances 0.000 claims description 6
- 229920000647 polyepoxide Polymers 0.000 claims description 6
- 238000002360 preparation method Methods 0.000 claims description 6
- 239000004917 carbon fiber Substances 0.000 claims description 5
- 229920001721 polyimide Polymers 0.000 claims description 5
- 229920000049 Carbon (fiber) Polymers 0.000 claims description 4
- 239000004699 Ultra-high molecular weight polyethylene Substances 0.000 claims description 4
- PNEYBMLMFCGWSK-UHFFFAOYSA-N aluminium oxide Inorganic materials [O-2].[O-2].[O-2].[Al+3].[Al+3] PNEYBMLMFCGWSK-UHFFFAOYSA-N 0.000 claims description 4
- 229920006231 aramid fiber Polymers 0.000 claims description 4
- XLJMAIOERFSOGZ-UHFFFAOYSA-M cyanate Chemical compound [O-]C#N XLJMAIOERFSOGZ-UHFFFAOYSA-M 0.000 claims description 4
- 238000010438 heat treatment Methods 0.000 claims description 4
- 239000011159 matrix material Substances 0.000 claims description 4
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims description 4
- 239000009719 polyimide resin Substances 0.000 claims description 4
- 229920005749 polyurethane resin Polymers 0.000 claims description 4
- 239000010453 quartz Substances 0.000 claims description 4
- VYPSYNLAJGMNEJ-UHFFFAOYSA-N silicon dioxide Inorganic materials O=[Si]=O VYPSYNLAJGMNEJ-UHFFFAOYSA-N 0.000 claims description 4
- 238000012546 transfer Methods 0.000 claims description 4
- 229920000785 ultra high molecular weight polyethylene Polymers 0.000 claims description 4
- 238000001816 cooling Methods 0.000 claims description 3
- 238000004519 manufacturing process Methods 0.000 abstract description 8
- 238000013461 design Methods 0.000 abstract description 5
- 230000010354 integration Effects 0.000 abstract 1
- 239000004593 Epoxy Substances 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 230000007704 transition Effects 0.000 description 2
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 239000004642 Polyimide Substances 0.000 description 1
- 238000013459 approach Methods 0.000 description 1
- 229910052799 carbon Inorganic materials 0.000 description 1
- 239000011248 coating agent Substances 0.000 description 1
- 238000000576 coating method Methods 0.000 description 1
- 238000005056 compaction Methods 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000005260 corrosion Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000001514 detection method Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000003733 fiber-reinforced composite Substances 0.000 description 1
- 239000003365 glass fiber Substances 0.000 description 1
- 239000002184 metal Substances 0.000 description 1
- 238000003825 pressing Methods 0.000 description 1
Abstract
The application provides a forming method of continuous layering of composite material blades of an aero-engine, and belongs to the technical field of composite materials of aero-engines. According to the method, the problem that the assembly and connection of the existing aero-engine composite material blade serving as an independent component are complex is solved, the composite material blade and the connecting component are designed into a whole to be paved through layering design, and then an integrated molded product is obtained through curing molding. The method simplifies the assembly process, solves the problem of solving assembly, and improves the integration performance and the production efficiency.
Description
Technical Field
The application belongs to the technical field of aeronautical composite materials, and particularly relates to a forming method for continuous layering of aeronautical engine composite material blades.
Background
Aeroengine composite blades are an important component in aeroengines, the main function of which is to convert a flow of gas into mechanical energy to drive an aircraft. Compared with the traditional metal blade, the composite material blade has better performance and efficiency due to the high strength, low density and good corrosion resistance. Aero-engine composite blades are typically made of fiber reinforced composites, wherein the fibers are typically carbon or glass fibers, and the matrix material is typically epoxy, polyimide, or the like. The choice of these materials may be determined according to the requirements of the blade and the working environment to achieve the desired strength, stiffness and durability.
In conventional manufacturing methods, aircraft engine composite blades are typically individually layered and cured and then assembled with a casing or the like. The composite material blade is used as an independent component for paving, the paving structure is simpler, and the problem of paving transition with other components is not considered. After the composite material blades are independently paved and solidified and formed, the composite material blades are required to be assembled and connected with parts such as a casing and the like through mechanical connection, the assembly process is complicated, and a large amount of manual operation and time are required; if the blades are densely distributed, it also presents a great challenge for independent assembly of the blades, as each blade needs to be precisely mounted in the correct position. In addition, because the blade and the casing are manufactured separately, there may be ply transition problems and joint strength concerns.
Disclosure of Invention
In view of the problem that the assembly and connection of the existing composite material blade of the aeroengine are complex as independent components, the application provides a forming method for continuously paving the composite material blade of the aeroengine.
In order to achieve the above purpose, the present application adopts the following technical scheme:
a molding method of continuous layering of composite material blades of an aeroengine comprises the following steps:
the preparation method of the skin prepreg comprises the following steps: performing layering of skin prepreg on the surface of a forming tool of the connecting part;
the preparation method of the blade prepreg comprises the following steps: layering a blade prepreg on the surface of a preformed tool of the blade, and performing preforming;
assembling the prepreg: transferring and placing the preformed 1 st blade prepreg on the skin prepreg surface of the connecting part; transferring and placing the preformed 2 nd blade prepreg on one side of the 1 st blade prepreg; repeating the assembly steps of the prepreg until the transfer and placement of all the blade prepregs are completed, and obtaining a composite material blade structure to be solidified;
and (3) curing and forming the blade: and (3) putting the composite material blade structure to be solidified into a vacuum bag, sending the vacuum bag into an autoclave for heating, pressurizing and solidifying, and taking out and cooling to obtain the composite material blade.
Preferably, the connecting member includes a case connecting member.
Preferably, the forming tool of the connecting part and the preforming tool of the blade are surface coated with a release material and dried before layering.
Preferably, the skin prepreg and the blade prepreg of the connecting component are respectively one of carbon fiber prepreg, quartz fiber prepreg, alumina fiber prepreg, aramid fiber prepreg and ultra-high molecular weight polyethylene fiber prepreg, and the resin matrix material is respectively one of epoxy resin, cyanate resin, polyurethane resin and polyimide resin.
Preferably, the lay-up of the skin prepreg of the connecting part is a symmetrical balanced lay-up, the thickness of the skin prepreg being determined from the stressed load.
Preferably, the layering of the blade prepreg is a symmetrical balanced layering, and the layering thickness is determined according to the stressed load of the blade.
Preferably, the method further comprises the step of preparing a blade reinforcing material: and (3) paving the blade reinforcing material on the surface of the semi-curing forming tool of the blade reinforcing material, and performing semi-curing forming.
Preferably, the assembly steps of the prepreg are: transferring and placing the preformed 1 st blade prepreg on the skin prepreg surface of the connecting part; placing the blade reinforcing material which is subjected to semi-curing molding on one side of the 1 st blade prepreg; transferring and placing the preformed 2 nd blade prepreg on a side of the blade reinforcement material facing away from the 1 st blade prepreg with the blade reinforcement material between the two blade prepregs; and repeating the assembly steps of the prepreg until the transfer and the placement of all the blade prepregs and the blade reinforcing materials are completed, and obtaining the composite material blade structure to be cured.
Preferably, the semi-cured molding of the blade reinforcing material is surface-coated with a release material and dried prior to layering.
Preferably, the prepreg used for the blade reinforcing material is the same as the blade prepreg, and one of carbon fiber prepreg, quartz fiber prepreg, alumina fiber prepreg, aramid fiber prepreg and ultra-high molecular weight polyethylene fiber prepreg is selected, and the resin matrix material is selected from one of epoxy resin, cyanate resin, polyurethane resin and polyimide resin.
Preferably, the blade reinforcing material is pressurized and cured by heating by using a hot press to realize semi-curing molding.
Compared with the traditional method, the application has the following beneficial effects:
1. simplifying the assembly process: in the traditional method, the composite material blade needs to be independently paved, solidified and formed, and then assembled and connected with parts such as a casing and the like. The method designs the composite material blade and the connecting part into a whole for layering through the whole layering design, so that the complicated steps caused by independent assembly of the composite material blade are eliminated. Thus, the assembly process can be greatly simplified, and the procedures and time are reduced.
2. Solves the assembly problem: in conventional approaches, if the blades are densely distributed, independent assembly can present a significant challenge. The continuous layering design in the method of the application ensures that the connection between the blades is more compact, can more easily solve the assembly problem of dense blades, and improves the feasibility and efficiency of assembly.
3. The integrated performance is improved: according to the method, through integral layering design and curing molding, the composite material blades and the connecting parts form an integrated molded product, 1 blade prepreg is a component part of 2 adjacent blades, the fiber prepregs among the surface blades are continuously layered, the blades are not independent individuals, and the structure strength is better and beneficial. Compared with the traditional independent assembly method, the integrated molding can provide stronger connection strength and structural consistency, and reduces the defects and weak points of material interfaces, thereby improving the overall strength and performance.
4. The production efficiency is improved: the method simplifies the manufacturing flow, reduces the independent assembly procedure and complexity, and saves the production time and the labor cost; the whole layering and integrated forming mode can also improve the automation degree of production and the efficiency of a production line, and further improve the production efficiency and the productivity.
Drawings
FIG. 1 is a schematic cross-sectional view of a continuous lay-up of a composite blade in an embodiment;
in the figure, 1: skin prepreg for connecting parts such as cases, 2: blade prepreg, 3: blade reinforcing material.
Fig. 2 is a schematic structural view of a preform tooling for a blade prepreg in an embodiment.
Fig. 3 is a schematic structural view of a preformed blade prepreg in an embodiment.
Fig. 4 is a schematic structural diagram of a semi-curing molding tool for a blade reinforcing material in an embodiment.
FIG. 5 is a schematic structural view of a pre-cured blade reinforcement material in an embodiment.
Detailed Description
In order to make the technical features and advantages or technical effects of the technical scheme of the application more obvious and understandable, the following detailed description is given with reference to the accompanying drawings.
The embodiment specifically discloses a method for forming continuous layering of composite material blades of an aeroengine, wherein a cross-section layering structure of composite material blade forming is shown in fig. 1, and the method comprises three parts of connecting part skin prepreg 1 such as a casing, blade prepreg 2 and blade reinforcing material 3, and the specific forming steps are as follows:
1) Coating a demolding material on the surfaces of all the molding tools, and respectively paving T300-level epoxy resin prepregs with the thickness of 2mm on the molding tools of connecting parts such as a casing and the like, the preforming tools of the blades (see figure 2) and the semi-curing molding tools of the blade reinforcing materials (see figure 4) after the molding tools are dried; preforming of the blade material is a room temperature pre-compaction process in the composite field, for example, the blade material comprises 10 layers of T300 grade epoxy resin prepregs, then the preforming process is: when the first layer of fiber prepreg is paved, pre-compacting is carried out for 10min at normal temperature, and the vacuum bag pressure is less than or equal to minus 0.09MPa; pre-compacting 3 layers of fiber prepreg at normal temperature for 10min, and vacuum bag pressing is less than or equal to-0.09 MPa; after the final layer of fiber prepreg is paved and pasted, the fiber prepreg is pre-compacted for 10 minutes at normal temperature, and the vacuum bag pressure is less than or equal to minus 0.09MPa.
2) Semi-curing and forming the laid blade reinforcing material, wherein the semi-curing condition is that the temperature is kept at 90 ℃ for 0.5h, and the semi-cured blade reinforcing material is shown in figure 5; the semi-curing molding process of the blade reinforcing material is in the field of composite materials, and if the blade reinforcing material adopts fiber prepreg of a medium-temperature curing epoxy system, the semi-curing process is to keep the temperature at 90 ℃ for 0.5-1 h, so that the resin in the blade reinforcing material has a certain curing degree.
3) Transferring the 1 st preformed blade prepreg (see figure 3) onto the skin prepreg of the connecting part such as the casing and the like, then placing the half-cured blade reinforcing material on one side of the blade prepreg, and then placing the 2 nd blade prepreg (see figure 3) on the other side of the blade reinforcing material, so as to ensure that the blade reinforcing material is between the 1 st and 2 nd blade prepregs;
4) In this example, there are 16 blade prepregs and 16 blade reinforcing materials, and the above operations are repeated until all the blade prepregs and the blade reinforcing materials are transferred and laid;
5) Preparing a vacuum bag, after the air tightness detection is qualified, sending the vacuum bag into an autoclave for heating, pressurizing and curing, wherein the curing system is that the molding pressure is 0.6MPa, and the temperature is kept for 2 hours at 130 ℃, and the curing molding conditions can be selected as required within the following ranges, namely, the molding pressure is 0.5-1.0 MPa, the temperature is kept at 120-150 ℃, and the time is 2-3 hours; and taking out and cooling to obtain the aeroengine composite material blade to be molded.
Example 2
This embodiment differs from embodiment 1 in that: the blade reinforcing material may be removed or replaced with another material. The effect of the blade reinforcement material of example 1 is to provide structural strength to the composite blade without affecting the continuous layering process of the composite blade. If the blade reinforcing material is replaced by a material with a certain structural strength and a wave absorbing function, other molding steps are unchanged, and the integrated composite material blade with the stealth function can be obtained.
Example 3
This embodiment differs from embodiment 1 in that: the number of the composite material blades is increased from 16 to 32, other forming steps are kept unchanged, and after the number of the blades is increased to 32, if a mode of single blade preparation and later assembly is adopted, the assembly workload is doubled, and the assembly difficulty is greatly increased. And the mode of integrally preparing the composite material blade is adopted, so that the workload of layering is only increased, and the preparation difficulty and the workload of the process are not increased much.
The technological conditions of preforming of the blade prepreg, semi-curing forming of the blade reinforcing material, final curing forming and the like are all conventional technical means in the field, and a person skilled in the art can select proper technological conditions according to the type of the prepreg.
Although the present application has been described with reference to the above embodiments, it should be understood that the application is not limited thereto, and that modifications and equivalents may be made thereto by those skilled in the art, which modifications and equivalents are intended to be included within the scope of the present application as defined by the appended claims.
Claims (10)
1. The forming method of the continuous layering of the composite material blade of the aeroengine is characterized by comprising the following steps of:
the preparation method of the skin prepreg comprises the following steps: performing layering of skin prepreg on the surface of a forming tool of the connecting part;
the preparation method of the blade prepreg comprises the following steps: layering a blade prepreg on the surface of a preformed tool of the blade, and performing preforming;
assembling the prepreg: transferring and placing the preformed 1 st blade prepreg on the skin prepreg surface of the connecting part; transferring and placing the preformed 2 nd blade prepreg on one side of the 1 st blade prepreg; repeating the assembly steps of the prepreg until the transfer and placement of all the blade prepregs are completed, and obtaining a composite material blade structure to be solidified;
and (3) curing and forming the blade: and (3) putting the composite material blade structure to be solidified into a vacuum bag, sending the vacuum bag into an autoclave for heating, pressurizing and solidifying, and taking out and cooling to obtain the composite material blade.
2. The molding method of claim 1, wherein the connecting member comprises a case connecting member.
3. The molding method as claimed in claim 1, wherein the molding tool of the connecting member and the preforming tool of the blade are surface-coated with a mold release material and dried before layering.
4. The molding method according to claim 1, wherein the skin prepreg and the blade prepreg of the connecting member are each one of carbon fiber prepreg, quartz fiber prepreg, alumina fiber prepreg, aramid fiber prepreg, and ultra-high molecular weight polyethylene fiber prepreg, and the resin base material is each one of epoxy resin, cyanate resin, polyurethane resin, and polyimide resin.
5. The molding method as claimed in claim 1, wherein the lay-up of the skin prepreg of the connecting part is a symmetrical balanced lay-up, and the thickness of the skin prepreg is determined from the load.
6. The method of claim 1, wherein the lay-up of the prepreg of the blade is a symmetrical balanced lay-up, and the thickness of the lay-up is determined based on the applied load of the blade.
7. The molding method as claimed in claim 1, further comprising a step of preparing a blade reinforcing material: layering the blade reinforcing material on the surface of the semi-curing forming tool of the blade reinforcing material, and performing semi-curing forming;
the prepreg is assembled by the following steps: transferring and placing the preformed 1 st blade prepreg on the skin prepreg surface of the connecting part; placing the blade reinforcing material which is subjected to semi-curing molding on one side of the 1 st blade prepreg; transferring and placing the preformed 2 nd blade prepreg on a side of the blade reinforcement material facing away from the 1 st blade prepreg with the blade reinforcement material between the two blade prepregs; and repeating the assembly steps of the prepreg until the transfer and the placement of all the blade prepregs and the blade reinforcing materials are completed, and obtaining the composite material blade structure to be cured.
8. The molding method as claimed in claim 7, wherein the semi-cured molding of the blade reinforcing material is performed before the laying, and the surface is coated with a mold release material and dried.
9. The molding method according to claim 7, wherein the prepreg used for the blade reinforcement material is the same as the blade prepreg, and one of carbon fiber prepreg, quartz fiber prepreg, alumina fiber prepreg, aramid fiber prepreg, and ultra-high molecular weight polyethylene fiber prepreg is selected, and the resin matrix material is one of epoxy resin, cyanate resin, polyurethane resin, and polyimide resin.
10. The molding method of claim 7, wherein the blade reinforcing material is pressed and heat-cured by a hot press to achieve semi-cured molding.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310983831.4A CN117162547A (en) | 2023-08-07 | 2023-08-07 | Forming method for continuous layering of composite material blade of aero-engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310983831.4A CN117162547A (en) | 2023-08-07 | 2023-08-07 | Forming method for continuous layering of composite material blade of aero-engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117162547A true CN117162547A (en) | 2023-12-05 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310983831.4A Pending CN117162547A (en) | 2023-08-07 | 2023-08-07 | Forming method for continuous layering of composite material blade of aero-engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117162547A (en) |
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2023
- 2023-08-07 CN CN202310983831.4A patent/CN117162547A/en active Pending
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