CN107266099B - Clamp for near-net forming of ceramic matrix composite turbine guide vane of aero-engine - Google Patents
Clamp for near-net forming of ceramic matrix composite turbine guide vane of aero-engine Download PDFInfo
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- CN107266099B CN107266099B CN201710457867.3A CN201710457867A CN107266099B CN 107266099 B CN107266099 B CN 107266099B CN 201710457867 A CN201710457867 A CN 201710457867A CN 107266099 B CN107266099 B CN 107266099B
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- C—CHEMISTRY; METALLURGY
- C04—CEMENTS; CONCRETE; ARTIFICIAL STONE; CERAMICS; REFRACTORIES
- C04B—LIME, MAGNESIA; SLAG; CEMENTS; COMPOSITIONS THEREOF, e.g. MORTARS, CONCRETE OR LIKE BUILDING MATERIALS; ARTIFICIAL STONE; CERAMICS; REFRACTORIES; TREATMENT OF NATURAL STONE
- C04B35/00—Shaped ceramic products characterised by their composition; Ceramics compositions; Processing powders of inorganic compounds preparatory to the manufacturing of ceramic products
- C04B35/71—Ceramic products containing macroscopic reinforcing agents
- C04B35/78—Ceramic products containing macroscopic reinforcing agents containing non-metallic materials
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Abstract
The invention discloses a clamp for near-net forming of a turbine guider blade made of an aeroengine ceramic matrix composite, wherein a cavity A (8) matched with a large blade preformed edge plate (3) is arranged at the top of an edge plate module (4), a cavity B (11) matched with a small edge plate (2) is arranged at the bottom of a small edge plate module (5), a curved surface (12) matched with a blade back airfoil (21) is arranged at the rear side of a blade back airfoil module (6), and a bulge (14) matched with a blade basin airfoil (22) is arranged at the front side of a blade basin airfoil module (7). The beneficial effects of the invention are as follows: the working surface of the finished blade is guaranteed to be free of or less in allowance for processing, fast to disassemble, assemble and assemble, near net forming of the ceramic matrix composite turbine guide blade is effectively achieved, and manufacturing cost is low.
Description
Technical Field
The invention relates to the technical field of manufacturing of aero-engine parts, in particular to a clamp for near-net forming of a ceramic matrix composite turbine guide vane of an aero-engine.
Background
The ceramic matrix composite turbine guide vane is used as one of the most critical parts of an aeroengine, and because the working environment of the turbine guide vane is very bad, the turbine guide vane has the highest temperature bearing, the highest strong hot corrosion bearing and the most serious thermal shock bearing, and the ceramic matrix composite has the excellent properties of small density, high specific strength, high specific stiffness, good high temperature resistance and the like, and is the ceramic matrix composite which is the most likely to replace the turbine guide vane high temperature material used by the nickel matrix superalloy at higher temperature. The ceramic matrix composite is generally carbon fiber reinforced carbon matrix (C/C) composite, carbon fiber reinforced silicon carbide ceramic matrix (C/SiC) composite and silicon carbide fiber reinforced silicon carbide ceramic matrix (SiC/SiC) composite, and is an ultrahigh temperature composite with the working temperature as high as 1650 ℃.
Currently, ceramic matrix composite turbine guide vanes are typically manufactured by braiding a vane preform, chemical vapor infiltration (chemical vapor infiltration, CVI) interfacial layer, matrix densification (using CVI and PIP techniques), machining, coating preparation, inspection and acceptance. The preparation of high performance ceramic matrix composite turbine guide vanes with three dimensional airfoil blade bodies, which must be woven into guide vane preforms by three dimensional weaving (e.g., 2.5D, three dimensional four-way, three dimensional five-way, orthogonal three-way, etc.), requires near net shape forming (i.e., no or little machining of the vane is allowed during preparation). However, the prefabricated body of the blade woven by the fibers is softer, and the design dimensions of the three-dimensional airfoil of the blade body and the like cannot be achieved by means of the woven blade. In the preparation processes of CVI deposition interface layer of the blade preform, the initial stage of matrix densification and the like, the shape and the size of the ceramic matrix composite material blade are difficult to ensure because the blade preform is not subjected to tooling, and defects such as warping, twisting, protruding and the like can occur, so that the preparation of the ceramic matrix composite material turbine guide blade fails. The structure of the blade preform is shown in fig. 1 and 2, the blade preform comprises a body (1), a small flange plate (2) and a large flange plate (3), the small flange plate (2) is fixedly connected to the top of the body (1), the small flange plate (2) is fixedly connected to the bottom of the body (1), the front part of the body (1) is provided with a blade back wing profile (21), and the rear part of the blade preform is provided with a blade basin wing profile (22).
Disclosure of Invention
The invention aims to overcome the defects of the prior art and provide the fixture for near-net forming of the ceramic matrix composite turbine guide vane of the aeroengine, which can ensure that the working surface of the finished vane has no or less allowance, is quickly disassembled, assembled and assembled, and effectively realizes near-net forming of the ceramic matrix composite turbine guide vane and has low manufacturing cost.
The aim of the invention is achieved by the following technical scheme: the fixture for near-net forming of the ceramic matrix composite turbine guide vane of the aeroengine comprises a large flange plate module, a small flange plate module, a vane back wing type module and a vane basin wing type module, wherein a cavity A matched with the large flange plate of a vane preform is arranged at the top of the large flange plate module, a vane back positioning block and a vane basin positioning block which are respectively positioned at the left side and the right side of the cavity A are also arranged on the large flange plate module, the vane back positioning block is positioned at the front side of the vane basin positioning block, and a cavity B matched with the small flange plate is arranged at the bottom of the small flange plate module; the back side of the back wing section module is provided with a curved surface matched with the back wing section, the bottom of the back wing section module is provided with a first positioning hole, the back wing section module is arranged on the flange plate module, and the first positioning hole is matched with the back positioning block; the front side of the leaf basin airfoil module is provided with a bulge matched with the leaf basin airfoil, the bottom of the leaf basin airfoil module is provided with a second positioning hole, the leaf basin airfoil module is arranged on the flange plate module, and the second positioning hole is matched with the leaf basin positioning block; the left side and the right side of the leaf basin wing type module and the left side of the leaf back wing type module are respectively and fixedly provided with a mounting plate, the leaf basin wing type module and the leaf back wing type module are respectively arranged between the large flange plate module and the small flange plate module, locking short screws are arranged between the mounting plates of the leaf basin wing type module and the mounting plates of the leaf back wing type module, locking long screws are arranged among the leaf basin wing type module, the large flange plate module and the small flange plate module, and locking long screws are arranged among the leaf back wing type module, the large flange plate module and the small flange plate module; the large flange plate module is characterized in that a plurality of small holes communicated with the cavity A are uniformly distributed on the outer surface of the large flange plate module, a plurality of small holes communicated with the cavity B are uniformly distributed on the outer surface of the small flange plate module, a plurality of small holes penetrating through a curved surface are distributed on the front surface of the back wing section module, and a plurality of small holes penetrating through a bulge are formed in the back of the back wing section module.
The two mounting plates on the airfoil module have the same structure.
Locking nuts are connected to the locking short screws and the locking long screws in a threaded mode.
The distance between two adjacent transverse small holes and the distance between two adjacent longitudinal small holes are 7-8 mm.
The diameter of the small hole is 2-3 mm.
The invention has the following advantages: (1) According to the invention, through adopting four modules designed by high-temperature resistant alloy or graphite materials and adopting the screwed screw and nut to apply force in a matched manner, the blade preform is fixed, compacted and compacted, and favorable conditions are provided for the near-net-size forming of the blade. (2) The fixture ensures that the working surface of the blade is processed without or with less allowance, avoids damaging the three-dimensional fiber woven structure of the blade, reduces the performance index of the blade, and improves the preparation quality and the production efficiency of the blade.
Drawings
FIG. 1 is a schematic view of a blade preform;
FIG. 2 is a right side view of FIG. 1;
FIG. 3 is a schematic diagram of the structure of the present invention;
FIG. 4 is a schematic view of the present invention with the back airfoil module and basin airfoil module removed;
FIG. 5 is a schematic view of the invention with the large edge plate module and the small edge plate module removed;
FIG. 6 is a schematic structural view of a large edge plate module;
FIG. 7 is a schematic view of a small flange module;
FIG. 8 is a schematic structural view of a blade back airfoil module;
FIG. 9 is a bottom view of FIG. 8;
FIG. 10 is a schematic structural view of a leaf basin airfoil module;
FIG. 11 is a bottom view of FIG. 10;
in the figure, a 1-body, a 2-small flange plate, a 3-large flange plate, a 4-large flange plate module, a 5-small flange plate module, a 6-back wing type module, a 7-leaf basin wing type module, an 8-cavity A, a 9-back positioning block, a 10-leaf basin positioning block, an 11-cavity B, a 12-curved surface, a 13-first positioning hole, a 14-bulge, a 15-second positioning hole, a 16-mounting plate, a 17-locking short screw, a 18-locking long screw, a 19-small hole, a 20-locking nut, a 21-back wing type and a 22-leaf basin wing type.
Detailed Description
The invention is further described below with reference to the accompanying drawings, the scope of the invention not being limited to the following:
as shown in fig. 3-11, the fixture for near-net forming of the ceramic matrix composite turbine guide vane of the aeroengine comprises a large flange plate module 4, a small flange plate module 5, a vane back airfoil module 6 and a vane basin airfoil module 7, wherein a cavity A8 matched with the large flange plate 3 of the vane preform is arranged at the top of the large flange plate module 4, a vane back positioning block 9 and a vane basin positioning block 10 which are respectively positioned at the left side and the right side of the cavity A8 are also arranged on the large flange plate module 4, the vane back positioning block 9 is positioned at the front side of the vane basin positioning block 10, and a cavity B11 matched with the small flange plate 2 is arranged at the bottom of the small flange plate module 5; the rear side of the back wing type module 6 is provided with a curved surface 12 matched with a back wing type 21, the bottom of the back wing type module 6 is provided with a first positioning hole 13, the back wing type module 6 is arranged on the flange plate module 4, and the first positioning hole 13 is matched with the back positioning block 9; the front side of the leaf basin airfoil module 7 is provided with a bulge 14 matched with a leaf basin airfoil 22, the bottom of the leaf basin airfoil module 7 is provided with a second positioning hole 15, the leaf basin airfoil module 7 is arranged on the flange plate module 4, and the second positioning hole 15 is matched with the leaf basin positioning block 10; the left side and the right side of the leaf basin airfoil module 7 and the leaf back airfoil module 6 are respectively and fixedly provided with a mounting plate 16, the leaf basin airfoil module 7 and the leaf back airfoil module 6 are respectively arranged between the large flange plate module 4 and the small flange plate module 5, a locking short screw 17 is arranged between the mounting plate 16 of the leaf basin airfoil module 7 and the mounting plate 16 of the leaf back airfoil module 6, a locking long screw 18 is arranged between the leaf basin airfoil module 7, the large flange plate module 4 and the small flange plate module 5, locking nuts 20 are respectively and fixedly connected with the locking short screw 17 and the locking long screw 18, and locking long screws 18 are arranged among the leaf back airfoil module 6, the large flange plate module 4 and the small flange plate module 5; the large flange plate module 4 is characterized in that a plurality of small holes 19 communicated with the cavity A8 are uniformly distributed on the outer surface of the large flange plate module 4, a plurality of small holes 19 communicated with the cavity B11 are uniformly distributed on the outer surface of the small flange plate module 5, a plurality of small holes 19 penetrating through the curved surface 12 are distributed on the front surface of the back wing type module 6, and a plurality of small holes 19 penetrating through the protrusions 14 are arranged on the rear surface of the back wing type module 7. The main function of the small hole is as follows: 1. in the densification process of the CVI preparation interface layer or the matrix, providing a channel to enable reaction gas to enter the blade preform for reaction and deposition of the ceramic matrix; 2. in PIP densification, channels are provided to allow the reaction solution to enter the blade preform for impregnation and then, during pyrolysis, to allow the reacted dead gas reactants to exit the blade preform.
The two mounting plates 16 on the airfoil module are identical in construction. The distance between two adjacent small holes 19 in the transverse direction and the distance between two adjacent small holes 19 in the longitudinal direction are 7-8 mm; the diameter of the small hole 19 is 2-3 mm.
The preparation of the near-net-shape of the ceramic matrix composite turbine guide vane of the aeroengine comprises the following steps:
s1, braiding a three-dimensional braiding turbine guide vane preform by adopting a special braiding machine, wherein the structure of the vane preform is shown in figures 1 and 2.
S2, placing the turbine guide vane preform in a vacuum high-temperature heat treatment furnace for high-temperature pretreatment, wherein the pretreatment temperature is 700-2400 ℃ and the time is 1-3 h, and protecting the turbine guide vane preform by adopting vacuum or inert gas, so that the pretreatment aims at removing residual glue on the surface of the vane preform and improving the physical performance of the vane preform;
s3, performing CVI interface layer on the turbine guide vane preform, wherein the CVI interface layer specifically comprises the following steps of:
s3 (I) placing the blade preform in the fixture, wherein the specific fixture process is as follows: the cavity B11 of the small flange plate module 5 is matched with the small flange plate 2 of the blade preform, so that the placing direction of the blade preform is ensured, the two are in smooth connection, and the problem of uneven height cannot occur; the cavity A8 of the large edge plate module 4 is matched with the large edge plate 3 of the blade preform, so that the placing direction of the blade preform is ensured, the two are in smooth connection, and the problem of uneven height cannot occur; assembling the blade back positioning blocks 9 on the large flange plate module 4, and assembling the blade back positioning blocks 9 into the first positioning holes 13 of the blade back airfoil modules 6 to determine the relative positions among the large flange plate module 4, the blade back airfoil modules 6 and the blade prefabricated body; the blade back wing type module 6, the small flange plate module 5 and the large flange plate module 4 are connected through locking long screws 18; assembling a leaf basin positioning block 10 on the large flange plate module 4, wherein the leaf basin positioning block 10 is assembled in a second positioning hole 15 of the leaf basin airfoil module 7 so as to determine the relative positions among the small flange plate module 5, the large flange plate module 4, the leaf basin airfoil module 7 and the blade preform; the leaf basin airfoil module 7, the small flange plate module 5 and the large flange plate module 4 are connected through locking long screws 18; the locking short screw 17 is connected with the locking nut 20 to fix the vane back wing type module 6 and the vane basin wing type module 7; tightening the locking nut 20 on the locking long screw 18 to compact and compress the blade preform, thereby realizing the quick tooling of the blade preform and ensuring that the molding size of the blade preform meets the design requirement of the drawing;
s3 (II) loading the fixture of the blade preform in the step S (I) into a CVI furnace, and adopting technical parameters to prepare a CVI interface layer, wherein the interface layer comprises a pyrolytic carbon interface layer and a BN interface layer. Methane or propane and natural gas are used as reaction gases in the preparation of pyrolytic carbon interface layer by CVI, and NH is used in the preparation of BN interface layer by CVI 3 Gas and BCl 3 The gas is used as reaction gas, the reaction gas is introduced into the clamp through the small holes 19 on the clamp, the CVI preparation temperature is 700-1100 ℃ and the CVI preparation time is 0.5-4 h.
S4, after the CVI is adopted to prepare the interface layer, the blade preform is still soft and not molded, and the matrix of the turbine guide blade preform needs to be densified in the initial stage. The method specifically comprises the following steps:
s4 (I) taking out the blade preform processed in the step S3 and loading the blade preform into another fixture, wherein the tooling mode is the same as that of the step S3 (I);
s4 (II) loading the fixture assembly with the blade preform into a CVI furnace or a PIP furnace, and adopting preparation technical parameters to perform ceramic matrix inductionDensification is produced, and the ceramic matrix comprises C and SiC. The CVI is used for preparing a C matrix by using methane or propane and natural gas as reaction gases, and the CVI is used for preparing a SiC matrix by using trichloromethylsilane (CH 3 SiCl 3 ）、H 2 Gas and BCl 3 The gas is used as reaction gas, the CVI preparation temperature is 700-1100 ℃ and the CVI preparation time is 20-100 h. Preparing a SiC matrix by precursor dipping and pyrolysis (precursor impregnation pyrolysis, PIP), dipping the SiC matrix by adopting polycarbosilane and xylene solution, then drying the SiC matrix in a baking oven, then performing high-temperature pyrolysis (pyrolysis temperature is 800-1200 ℃ and time is 0.5-3 h) in a vacuum high-temperature treatment furnace, and then repeatedly circulating the dipping, drying and pyrolysis processes for 3-7 times;
s4 (III) taking the clamp in the step S4 (2) out of the CVI furnace or the PIP furnace;
s4 (IV) taking out the blade preform from the clamp, wherein the taking-out process is the reverse process of the tooling process, checking whether the blade preform is hardened and shaped after taking out, if so, sending the blade preform into the next process, and if not, repeating the step S4 (II);
s5, final matrix densification of the turbine guide blade preform, wherein the density of the turbine guide blade preform reaches the final design requirement (the required density is 1.8g/cm 3 ～2.6ｇ/cm 3 ) The method comprises the following specific operation steps:
s5 (I), placing the blade preform into a CVI furnace or a PIP furnace for final ceramic matrix densification preparation, wherein the final ceramic matrix densification of the blade preform is a complementary process for matrix densification preparation in an initial stage, and the preparation conditions of reaction gas, preparation parameters, preparation process and the like are the same, but the preparation time of the blade preform is longer, for example, the preparation time of C matrix or SiC matrix by CVI is 200-600 h; or if PIP is adopted to prepare the SiC matrix, the repeated cycle times are 9-13 times.
S5 (II) taking the blade preform out of the CVI furnace or the PIP furnace, wherein the density of the blade preform of the turbine guider reaches the final required density, the substrate is completely densified, the substrate is completely hardened and shaped, and the blade preform of the turbine guider is made into a ceramic matrix composite material blade blank;
s6, machining a ceramic matrix composite blade blank. Because the working face size of the ceramic matrix composite blade blank in step S5 (II) is nearly net-shaped, other non-working face sizes must be machined according to design requirements.
S7, preparing a coating of the ceramic matrix composite blade. In order to improve the service performance of the ceramic matrix composite blade in the working environment of the aeroengine such as high-temperature oxidation, high-temperature corrosion and the like, a coating is required to be prepared on the surface of the ceramic matrix composite blade, for example, a chemical vapor deposition (chemical vapor deposition, CVD) method is adopted to prepare a SiC coating: and placing the blade in a CVD furnace, and carrying out CVD by adopting corresponding preparation technical parameters to prepare the SiC coating. CVD preparation of SiC matrix Using Trichloromethylsilane (CH) 3 SiCl 3 ）、H 2 The gas is taken as reaction gas, the gas enters through the small holes 19, the CVD preparation temperature is 900-1200 ℃ and the time is 10-60 h;
s8, checking and accepting the ceramic matrix composite turbine guide vane. After the preparation process is finished, all preparation processes of the ceramic matrix composite turbine guide vane are required to be checked and accepted according to the requirements of vane design drawing, and the qualified ceramic matrix composite turbine guide vane finished product is obtained after passing the process.
Because the top of the large edge plate module 4 is provided with the cavity A8 matched with the large edge plate 3 of the blade preform, the bottom of the small edge plate module 5 is provided with the cavity B11 matched with the small edge plate 2, the rear side of the blade back airfoil module 6 is provided with the curved surface 12 matched with the blade back airfoil 21, the front side of the blade basin airfoil module 7 is provided with the bulge 14 matched with the blade basin airfoil 22, and when the near net forming is carried out, the working surface of the blade is ensured to be processed without or with less allowance, the three-dimensional fiber weaving structure of the blade is prevented from being damaged, the performance index of the blade is prevented from being reduced, and the preparation quality and the production efficiency of the blade are improved. In addition, in the fixture, the leaf basin wing type module 7 and the leaf back wing type module 6 are locked through the matching of the locking short screw 17 and the locking nut 20, the leaf basin wing type module 7, the large flange plate module 4 and the small flange plate module 5 are locked through the matching of the locking long screw 18 and the locking nut 20, the leaf back wing type module 6, the large flange plate module 4 and the small flange plate module 5 are locked through the matching of the locking long screw 18 and the locking nut 20, the screwed screws and the nuts are used for matching force application, and the blade prefabricated body is fixed, compressed and compacted, so that favorable conditions are provided for the near net size forming of the blade, and the fixture is very convenient to use in the process of assembly and disassembly, and the production efficiency is saved and improved.
Claims (5)
1. The application method of the fixture for near-net forming of the ceramic matrix composite turbine guide vane of the aeroengine is characterized by comprising the following steps of: the fixture comprises a large flange plate module (4), a small flange plate module (5), a blade back wing type module (6) and a blade basin wing type module (7), wherein a cavity A (8) matched with a large flange plate (3) of a blade prefabricated body is arranged at the top of the large flange plate module (4), a blade back positioning block (9) and a blade basin positioning block (10) which are respectively positioned at the left side and the right side of the cavity A (8) are further arranged on the large flange plate module (4), the blade back positioning block (9) is positioned at the front side of the blade basin positioning block (10), and a cavity B (11) matched with the small flange plate (2) is arranged at the bottom of the small flange plate module (5); the back side of the back wing section module (6) is provided with a curved surface (12) matched with the back wing section (21), the bottom of the back wing section module (6) is provided with a first positioning hole (13), the back wing section module (6) is placed on the large flange plate module (4), and the first positioning hole (13) is matched with the back positioning block (9); the front side of the leaf basin airfoil module (7) is provided with a bulge (14) matched with a leaf basin airfoil (22), the bottom of the leaf basin airfoil module (7) is provided with a second positioning hole (15), the leaf basin airfoil module (7) is placed on the large flange plate module (4), and the second positioning hole (15) is matched with the leaf basin positioning block (10); the left side and the right side of the leaf basin airfoil module (7) and the left side of the leaf back airfoil module (6) are fixedly provided with mounting plates (16), the leaf basin airfoil module (7) and the leaf back airfoil module (6) are arranged between the large edge plate module (4) and the small edge plate module (5), locking short screws (17) are arranged between the mounting plates (16) of the leaf basin airfoil module (7) and the mounting plates (16) of the leaf back airfoil module (6), locking long screws (18) are arranged among the leaf basin airfoil module (7), the large edge plate module (4) and the small edge plate module (5), and locking long screws (18) are arranged among the leaf back airfoil module (6), the large edge plate module (4) and the small edge plate module (5); the large flange plate module (4) is characterized in that a plurality of small holes (19) communicated with the cavity A (8) are uniformly distributed on the outer surface of the large flange plate module (4), a plurality of small holes (19) communicated with the cavity B (11) are uniformly distributed on the outer surface of the small flange plate module (5), a plurality of small holes (19) penetrating through the curved surface (12) are distributed on the front surface of the blade back airfoil module (6), and a plurality of small holes (19) penetrating through the protrusions (14) are formed in the rear surface of the blade basin airfoil module (7);
the using method of the fixture comprises the steps of placing the blade preform in the fixture, wherein the specific tooling process is that,
the cavity B (11) of the small flange plate module (5) is matched with the small flange plate (2) of the blade preform, so that the placing direction of the blade preform is ensured, the two are in smooth connection, and the problem of uneven height cannot occur;
the cavity A (8) of the large edge plate module (4) is matched with the large edge plate (3) of the blade preform, so that the placing direction of the blade preform is ensured, the two are in smooth connection, and the problem of uneven height cannot occur;
the method comprises the steps that a blade back positioning block (9) is assembled on a large flange plate module (4), and the blade back positioning block (9) is assembled in a first positioning hole (13) of a blade back wing type module (6);
the blade back wing type module (6), the small flange plate module (5) and the large flange plate module (4) are connected through locking long screws (18);
the leaf basin positioning block (10) is assembled on the large flange plate module (4), and the leaf basin positioning block (10) is assembled in a second positioning hole (15) of the leaf basin airfoil module (7);
the leaf basin airfoil module (7), the small flange plate module (5) and the large flange plate module (4) are connected through locking long screws (18);
a locking short screw (17) is connected with a locking nut (20) to fix the blade back wing type module (6) and the blade basin wing type module (7); a lock nut (20) on the lock long screw (18) is screwed down to compact and press the blade preform.
2. The method of using the fixture for near-net-shape forming of an aero-engine ceramic matrix composite turbine guide vane as claimed in claim 1, wherein: the two mounting plates (16) on the airfoil module are identical in structure.
3. The method of using the fixture for near-net-shape forming of an aero-engine ceramic matrix composite turbine guide vane as claimed in claim 1, wherein: the locking short screw (17) and the locking long screw (18) are both in threaded connection with a locking nut (20).
4. The method of using the fixture for near-net-shape forming of an aero-engine ceramic matrix composite turbine guide vane as claimed in claim 1, wherein: the transverse distance and the longitudinal distance between the two small holes (19) are 7-8 mm.
5. The method for using the fixture for near-net-shape forming of the ceramic matrix composite turbine guide vane of the aeroengine as claimed in claim 4, wherein the method comprises the following steps: the diameter of the small hole (19) is 2-3 mm.
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| Application Number | Priority Date | Filing Date | Title |
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| CN201710457867.3A CN107266099B (en) | 2017-06-16 | 2017-06-16 | Clamp for near-net forming of ceramic matrix composite turbine guide vane of aero-engine |
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| Application Number | Priority Date | Filing Date | Title |
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| CN201710457867.3A CN107266099B (en) | 2017-06-16 | 2017-06-16 | Clamp for near-net forming of ceramic matrix composite turbine guide vane of aero-engine |
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| CN111102017B (en) * | 2019-12-13 | 2022-07-12 | 西安鑫垚陶瓷复合材料有限公司 | Aeroengine ceramic matrix composite fixed guider blade structure and forming method thereof |
| CN112552068B (en) * | 2020-11-26 | 2022-12-20 | 西安鑫垚陶瓷复合材料有限公司 | Protection tool and method for rivet welding of C/SiC-ZrC and C/SiC ceramic matrix composite |
| CN113733596B (en) * | 2021-08-30 | 2022-08-19 | 北京航空航天大学 | Composite material light undercarriage wheel structure and forming mode thereof |
| CN115093231B (en) * | 2022-06-23 | 2023-09-01 | 西安鑫垚陶瓷复合材料有限公司 | Ceramic matrix composite guide vane with trailing edge split joint and preparation method thereof |
| US20240229239A9 (en) * | 2022-10-21 | 2024-07-11 | Raytheon Technologies Corporation | Floating tooling assembly for chemical vapor infiltration |
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