CN113929482A - Ceramic matrix composite turbine guide vane and preparation method thereof - Google Patents
Ceramic matrix composite turbine guide vane and preparation method thereof Download PDFInfo
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
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- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
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- C04B2235/02—Composition of constituents of the starting material or of secondary phases of the final product
- C04B2235/30—Constituents and secondary phases not being of a fibrous nature
- C04B2235/38—Non-oxide ceramic constituents or additives
- C04B2235/3852—Nitrides, e.g. oxynitrides, carbonitrides, oxycarbonitrides, lithium nitride, magnesium nitride
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Abstract
The invention discloses a ceramic matrix composite turbine guide vane and a preparation method thereof. The invention also discloses a preparation method of the blade, which comprises the following steps: firstly preparing an inner mold, preparing a blade body prefabricated body according to the inner mold, then performing incision treatment, then preparing a plate-shaped prefabricated body, then penetrating the plate-shaped prefabricated body into the blade body prefabricated body, then performing flanging, sewing and shaping treatment, then sequentially depositing an interface layer and a silicon carbide ceramic matrix on the surface of the prefabricated body, and after reprocessing to the designed size, performing damage repair to obtain the ceramic matrix composite turbine guide blade. The method avoids splicing of the structural units by integrally forming the blade body and the edge plate, improves the integrity and the structural reliability of the member, has wide application range and can provide support for batch production.
Description
Technical Field
The invention relates to the technical field of gas turbine engine manufacturing, in particular to a ceramic matrix composite turbine guide vane and a preparation method thereof.
Background
In the structures of gas turbine engines such as aeroengines, gas turbines and the like, a turbine system is used for converting partial heat energy and potential energy in high-temperature gas into mechanical work and driving a gas compressor and accessories to work, the turbine system is a system with the largest heat load and power load in the gas turbine engine, and is characterized by large output power, high use temperature, light weight requirement and small structural size, a guide blade at the inlet of a high-pressure turbine is a part with the highest working temperature in the turbine system, and the turbine system is used for converting partial heat energy of high-temperature gas flow into kinetic energy and simultaneously enabling the gas flow to flow out in a certain direction so as to meet the gas flow and inlet direction required by the working turbine. At present, high-temperature alloy materials commonly used for the turbine guide blade of an engine have the problems of heat resistance temperature of not higher than 1100 ℃, heavy weight and the like, and carbon fiber reinforced carbon matrix composite materials capable of resisting higher temperature have the defect of high temperature and easy oxidation. The density of the ceramic matrix composite is only 1/3-1/4 of high-temperature alloy, the heat-resistant temperature is 150-350 ℃ higher than that of the high-temperature alloy, the ceramic matrix composite is resistant to acid and alkali corrosion and high in toughness, and meanwhile, an oxide protective film generated by the reaction of the ceramic matrix composite in a high-temperature gas environment can block cracks and pores on the surface of the ceramic matrix composite and prevent external oxygen from diffusing into the ceramic matrix composite, so that the high-temperature stability and the long service life of a component are ensured, and therefore the ceramic matrix composite is considered as one of the first-choice materials of a new-generation aircraft engine thermal protection component at home and abroad.
When the ceramic matrix composite material is used as a main structure material for designing the guide vane, a split structure scheme is generally adopted, namely a vane body and a flange plate of the vane are respectively prepared, and then the vane is integrally assembled through splicing, assembling, post-deposition and other modes.
Disclosure of Invention
In order to solve the technical problems, the invention aims to provide a ceramic matrix composite turbine guide vane and a preparation method thereof, so as to solve the problem that the connection reliability between a vane body and a flange plate has a large risk when the turbine guide vane is prepared in the prior art.
The technical scheme for solving the technical problems is as follows: the turbine guide vane is made of ceramic matrix composite materials, and a vane body and a flange plate of the turbine guide vane are integrally molded.
The invention has the beneficial effects that: the blade body and the flange plate of the turbine guide blade are integrally formed, so that the splicing between the structural units is avoided, and the connection reliability between the blade body and the flange plate is greatly enhanced.
On the basis of the technical scheme, the invention can be further improved as follows:
furthermore, the reinforcement of the ceramic matrix composite material is carbon fiber and/or silicon carbide fiber, and the matrix is silicon carbide.
The invention also provides a preparation method of the ceramic matrix composite turbine guide vane, which comprises the following steps:
(1) taking the molded surface of the inner cavity of the turbine guide blade as reference, and preparing an inner mold from graphite, wherein the inner mold is provided with a plurality of vent holes vertical to the molded surface;
(2) winding the woven fiber cloth on the outer surface of the inner mold, then sewing by using a sewing line made of the same material as the woven fiber cloth, wherein the wound and woven fiber cloth after sewing is a prefabricated body of the blade body, and then performing incision treatment at positions 5-10mm away from two ends of the prefabricated body of the blade body respectively, wherein the incision direction is parallel to the longitudinal direction of the prefabricated body of the blade body, and the incision length is 10-50 mm;
(3) preparing 4 same plate-shaped prefabricated bodies by using woven fiber cloth made of the same material as the prefabricated bodies of the blade bodies, and forming holes at corresponding positions of the blade bodies, wherein the shapes of the holes are the same as the sections of the prefabricated bodies of the blade bodies;
(4) penetrating 4 sheet-shaped prefabricated bodies into the blade body prefabricated body along the opening, wherein 2 sheets are positioned on two sides of the upper notch of the blade body prefabricated body, the other 2 sheets are positioned on two sides of the lower notch of the blade body prefabricated body, then respectively turning out the woven fiber cloth at the notch of the blade body prefabricated body along the outer surface of the blade body, clamping the turned edges between the sheet-shaped prefabricated bodies, respectively sewing the 2 sheet-shaped prefabricated bodies and the turned edges clamped in the middle into a whole by using a suture thread made of the same material as the blade body prefabricated body, and finally fixing the blade body prefabricated body and the sheet-shaped prefabricated bodies by using an inner mold to finish shaping;
(5) sequentially depositing an interface layer and a silicon carbide ceramic matrix on the surfaces of the shaped blade body prefabricated body and the shaped plate-shaped prefabricated body, and removing the inner mold to obtain a ceramic matrix composite turbine guide blade blank;
(6) and processing the ceramic matrix composite turbine guide vane blank to a design size, then continuously depositing silicon carbide on the surface of the ceramic matrix composite turbine guide vane blank, and performing damage repair to obtain the ceramic matrix composite turbine guide vane.
Further, the diameter of the vent hole in the step (1) is 2-5 mm.
Further, the graphite in the step (1) is electrode graphite or high-purity graphite.
Further, the raw material of the woven fiber cloth in the step (2) is carbon fiber and/or silicon carbide fiber.
Further, the weaving method in the step (2) is two-dimensional plain weaving, two-dimensional satin weaving, two-dimensional twill weaving or 2.5-dimensional weaving.
Further, in the step (2), the distance between the notches is 10-13mm.
Further, the thickness of the plate-shaped preform in the step (3) is 0.3 to 0.5 times of the design thickness of the flange plate.
Further, the step (5) of depositing the interface layer and the silicon carbide ceramic matrix is depositing by chemical vapor deposition.
Further, the interface layer in the step (5) is a boron nitride interface layer.
Further, the preparation process of the boron nitride interface layer is as follows: heating to 650-1000 Pa, keeping the temperature for 1-2h, introducing a mixed gas of argon, hydrogen, ammonia and boron trichloride, depositing for 15-35h, keeping the temperature for 1-2h, and cooling to room temperature; wherein the flow ratio of the argon gas, the hydrogen gas, the ammonia gas and the boron trichloride gas is 1: 1-3: 2-8: 2-8.
Further, the preparation process of the boron nitride interface layer is performed 1-3 times.
Further, the preparation process of the silicon carbide ceramic matrix comprises the following steps: under the condition of pressure of 200-5000Pa, heating to 900-1200 ℃, keeping the temperature for 1-2h, introducing mixed gas of trichloromethylsilane, hydrogen and argon, depositing for 30-80h, keeping the temperature for 2h, and cooling to room temperature; this preparation process was performed 4-8 times.
Further, the flow ratio of trichloromethylsilane, hydrogen and argon is 1: 5-15: 10-20.
Further, the processing in the step (6) is processing by using a machine or a laser.
Further, the process of depositing the silicon carbide in the step (6) is as follows: under the condition of pressure of 200-5000Pa, heating to 900-1200 ℃, keeping the temperature for 1-2h, introducing mixed gas of trichloromethylsilane, hydrogen and argon, depositing for 30-80h, keeping the temperature for 2h, and cooling to room temperature; this preparation process was performed 1-3 times.
The invention has the following beneficial effects:
according to the invention, by utilizing the flexibility characteristic of the fiber prefabricated body, the flanging structures at two ends of the blade body are obtained in a mode of in-situ grooving on the blade body prefabricated body, and are integrally sewn and formed with the flange plate prefabricated body, and synchronous densification is completed in the subsequent process. In addition, the reinforcing effect of the sewing fibers can improve the connection strength between the blade body and the flange plate by about 20-35%, and the structural reliability of the member is improved.
The method disclosed by the invention can be suitable for preparing single-connection and multi-connection guide vanes, is a near-net-size preparation method, has a wide application range, and can provide support for batch production of the ceramic matrix composite turbine guide vanes.
Drawings
FIG. 1 is a schematic cut-out view of a body preform according to the present invention;
FIG. 2 is a schematic view of a plate-shaped preform according to the present invention;
FIG. 3 is a schematic view of the flanging of the blade body preform according to the present invention;
FIG. 4 is a schematic view showing the shaping of a blade body preform and a plate-like preform in example 1;
FIG. 5 is a photograph of a ceramic matrix composite turbine vane prepared in example 1.
Wherein, 1, an inner mold; 2. a blade body preform; 3. cutting; 4. a plate-shaped preform; 5. opening a hole; 6. and (5) flanging.
Detailed Description
The principles and features of this invention are described below in conjunction with the following drawings, which are set forth by way of illustration only and are not intended to limit the scope of the invention. The examples, in which specific conditions are not specified, were conducted under conventional conditions or conditions recommended by the manufacturer. The reagents or instruments used are not indicated by the manufacturer, and are all conventional products available commercially.
Example 1:
the preparation method of the ceramic matrix composite turbine guide vane comprises the following steps:
(1) taking the molded surface of the inner cavity of the turbine guide blade as reference, preparing an inner mold 1 by using electrode graphite, wherein the inner mold 1 is provided with a plurality of vent holes which are vertical to the molded surface and have the diameter of 3 mm;
(2) winding two-dimensional plain woven carbon fiber cloth on the outer surface of an inner mold 1, sewing by using a carbon fiber sewing thread, wherein the sewn wound carbon fiber cloth is the blade body prefabricated body 2, and performing incision treatment at positions 8mm away from two ends on the blade body prefabricated body 2 respectively, wherein the direction of an incision 3 is parallel to the longitudinal direction of the blade body prefabricated body 2, the length of the incision 3 is 40mm, and the distance between the incisions 3 is 12 mm;
(3) the two-dimensional plain woven carbon fiber cloth is sewn in a laminated mode, 4 same plate-shaped prefabricated bodies 4 are prepared, the thickness of each plate is 0.35 times of the design thickness of the edge plate, then, holes are formed in the corresponding blade body positions in a cutting mode, and the shape of each hole 5 is the same as the cross section of each blade body;
(4) penetrating 4 sheets of plate-shaped prefabricated bodies 4 into the blade body prefabricated body 2 along the opening 5 in the step (3), wherein 2 sheets are positioned at two sides of the upper notch 3 of the blade body prefabricated body 2, the other 2 sheets are positioned at two sides of the lower notch 3 of the blade body prefabricated body 2, turning out the two-dimensional plain woven carbon fiber cloth at the notch 3 of the blade body prefabricated body 2 along the outer surface of the blade body, clamping the turned-over edges 6 between the plate-shaped prefabricated bodies 4, and sewing the 2 sheets of plate-shaped prefabricated bodies 4 and the turned-over edges 6 clamped in the middle into a whole by using carbon fiber sewing threads; fixing the blade body prefabricated body 2 and the plate-shaped prefabricated body 4 by using the inner mold 1 to finish shaping;
(5) putting the shaped blade body preform 2, the plate-shaped preform 4 and the inner mold 1 into a chemical vapor deposition furnace together, sequentially depositing a boron nitride interface layer and a silicon carbide ceramic matrix on the surface of the preform, and removing the inner mold 1 to obtain a ceramic matrix composite turbine guide blade blank;
the preparation process of the boron nitride interface layer comprises the following steps: heating to 680 ℃ under the condition that the pressure is 550Pa, preserving heat for 2 hours, introducing mixed gas of argon, hydrogen, ammonia and boron trichloride gas, wherein the flow ratio of the argon, the hydrogen, the ammonia and the boron trichloride gas is 1: 1-3: 2-8: 2-8, after depositing for 30h, continuing to preserve heat for 2h, and cooling to room temperature; this step is performed cyclically 2 times;
the preparation process of the silicon carbide ceramic matrix comprises the following steps: heating to 1050 ℃ under the condition that the pressure is 1200a, keeping the temperature for 2h, introducing mixed gas of trichloromethylsilane, hydrogen and argon, wherein the flow ratio of trichloromethylsilane to hydrogen to argon is 1: 12: 15, after depositing for 72 hours, continuing to preserve heat for 2 hours, and cooling to room temperature; this step is performed in a loop 7 times;
(6) machining a ceramic matrix composite turbine guide vane blank to a design size by using machinery or laser to obtain a semi-finished product of the ceramic matrix composite turbine guide vane, then placing the semi-finished product in a silicon carbide chemical vapor deposition furnace, depositing silicon carbide ceramic on all machining surfaces of the ceramic matrix composite turbine guide vane semi-finished product (the preparation process is the same as that of the silicon carbide ceramic matrix in the step (5)), repairing damage, and performing cyclic deposition for 2 times to obtain the ceramic matrix composite turbine guide vane.
Example 2:
the preparation method of the ceramic matrix composite turbine guide vane comprises the following steps:
(1) taking the molded surface of the inner cavity of the turbine guide blade as reference, preparing an inner mold 1 by using electrode graphite, wherein the inner mold 1 is provided with a plurality of vent holes which are vertical to the molded surface and have the diameter of 2 mm;
(2) winding two-dimensional satin weaving silicon carbide fiber cloth on the outer surface of an inner mold 1, then sewing by using a silicon carbide fiber sewing line, wherein the sewed silicon carbide fiber cloth is a blade body prefabricated body 2, and then performing incision treatment at positions 5mm away from two ends on the blade body prefabricated body 2 respectively, wherein the direction of an incision 3 is parallel to the longitudinal direction of the blade body prefabricated body 2, the length of the incision 3 is 10mm, and the distance between the incisions 3 is 10 mm;
(3) the two-dimensional satin weaving silicon carbide fiber cloth is overlapped and sewn, 4 same plate-shaped prefabricated bodies 4 are prepared, the thickness of each plate is 0.3 time of the design thickness of the edge plate, then, holes are formed in the corresponding blade body positions in a cutting mode, and the shape of each hole 5 is the same as the cross section of the blade body;
(4) penetrating 4 sheets of plate-shaped prefabricated bodies 4 into the blade body prefabricated body 2 along the opening 5 in the step (3), wherein 2 sheets are positioned at two sides of the upper notch 3 of the blade body prefabricated body 2, the other 2 sheets are positioned at two sides of the lower notch 3 of the blade body prefabricated body 2, turning out the two-dimensional satin weaving silicon carbide fiber cloth at the notch 3 of the blade body prefabricated body 2 along the outer surface of the blade body, clamping the flanging 6 between the plate-shaped prefabricated bodies 4, and sewing the 2 sheets of plate-shaped prefabricated bodies 4 and the flanging 6 clamped in the middle into a whole by using a silicon carbide fiber suture line; fixing the blade body prefabricated body 2 and the plate-shaped prefabricated body 4 by using the inner mold 1 to finish shaping;
(5) putting the shaped blade body preform 2, the plate-shaped preform 4 and the inner mold 1 into a chemical vapor deposition furnace together, sequentially depositing a boron nitride interface layer and a silicon carbide ceramic matrix on the surface of the preform, and removing the inner mold 1 to obtain a ceramic matrix composite turbine guide blade blank;
the preparation process of the boron nitride interface layer comprises the following steps: heating to 650 ℃ under the condition that the pressure is 50Pa, preserving heat for 2 hours, and then sequentially introducing argon, hydrogen, ammonia and boron trichloride gas, wherein the flow ratio of the argon to the hydrogen to the ammonia to the boron trichloride gas is 1: 1-3: 2-8: 2-8, after depositing for 35h, continuing to preserve heat for 2h, and cooling to room temperature; this step is performed cyclically 3 times;
the preparation process of the silicon carbide ceramic matrix comprises the following steps: heating to 900 ℃ under the pressure of 200Pa, keeping the temperature for 2h, and introducing mixed gas of trichloromethylsilane, hydrogen and argon, wherein the flow ratio of trichloromethylsilane to hydrogen to argon is 1: 5: 10, after depositing for 80 hours, continuing to preserve heat for 2 hours, and cooling to room temperature; this step is performed cyclically 8 times;
(6) machining the ceramic matrix composite turbine guide vane blank to a design size by using machinery or laser to obtain a semi-finished product of the ceramic matrix composite turbine guide vane, then placing the semi-finished product in a silicon carbide chemical vapor deposition furnace, depositing silicon carbide ceramic on all machining surfaces of the ceramic matrix composite turbine guide vane semi-finished product (the preparation process is the same as that of the silicon carbide ceramic matrix in the step (5)), repairing damage, and performing cyclic deposition for 3 times to obtain the ceramic matrix composite turbine guide vane.
Example 3:
the preparation method of the ceramic matrix composite turbine guide vane comprises the following steps:
(1) taking the molded surface of the inner cavity of the turbine guide blade as reference, preparing an inner mold 1 by using electrode graphite, wherein the inner mold 1 is provided with a plurality of vent holes which are vertical to the molded surface and have the diameter of 5 mm;
(2) winding two-dimensional twill woven carbon fiber cloth on the outer surface of an inner mold 1, then sewing by using a carbon fiber sewing line, wherein the wound carbon fiber cloth after sewing is the blade body prefabricated body 2, and then performing incision treatment at positions 10mm away from two ends on the blade body prefabricated body 2 respectively, wherein the direction of an incision 3 is parallel to the longitudinal direction of the blade body prefabricated body 2, the length of the incision 3 is 50mm, and the distance between the incisions 3 is 13 mm;
(3) the two-dimensional twill woven carbon fiber cloth is overlapped and sewn, 4 same plate-shaped prefabricated bodies 4 are prepared, the thickness of each plate is 0.5 time of the design thickness of a flange plate, then, holes are formed in the corresponding blade body positions in a cutting mode, and the shape of each hole 5 is the same as the cross section of each blade body;
(4) penetrating 4 sheets of plate-shaped prefabricated bodies 4 into the blade body prefabricated body 2 along the opening 5 in the step (3), wherein 2 sheets are positioned at two sides of the upper notch 3 of the blade body prefabricated body 2, the other 2 sheets are positioned at two sides of the lower notch 3 of the blade body prefabricated body 2, then respectively turning the woven fiber cloth at the notch 3 of the blade body prefabricated body 2 out along the outer surface of the blade body, clamping the turned edges 6 between the plate-shaped prefabricated bodies 4, and respectively sewing the 2 sheets of plate-shaped prefabricated bodies 4 and the turned edges 6 clamped in the middle into a whole by using carbon fiber sewing threads; fixing the blade body prefabricated body 2 and the plate-shaped prefabricated body 4 by using the inner mold 1 to finish shaping;
(5) putting the shaped blade body preform 2, the plate-shaped preform 4 and the inner mold 1 into a chemical vapor deposition furnace together, sequentially depositing a boron nitride interface layer and a silicon carbide ceramic matrix on the surface of the preform, and removing the inner mold 1 to obtain a ceramic matrix composite turbine guide blade blank;
the preparation process of the boron nitride interface layer comprises the following steps: heating to 1000 ℃ under the condition that the pressure is 1000Pa, preserving heat for 1h, and then sequentially introducing argon, hydrogen, ammonia and boron trichloride gas, wherein the flow ratio of the argon to the hydrogen to the ammonia to the boron trichloride gas is 1: 1-3: 2-8: 2-8, after depositing for 15h, continuing to preserve heat for 1h, and cooling to room temperature;
the preparation process of the silicon carbide ceramic matrix comprises the following steps: heating to 1200 ℃ under the condition that the pressure is 5000Pa, preserving heat for 1h, introducing mixed gas of trichloromethylsilane, hydrogen and argon, wherein the flow ratio of trichloromethylsilane to hydrogen to argon is 1: 15: 20, after depositing for 30 hours, continuing to preserve heat for 1 hour, and cooling to room temperature; this step is performed cyclically 4 times;
(6) and (4) machining the ceramic matrix composite turbine guide vane blank to a designed size by using machinery or laser to obtain a semi-finished product of the ceramic matrix composite turbine guide vane, then placing the semi-finished product in a silicon carbide chemical vapor deposition furnace, depositing silicon carbide ceramic on all machining surfaces of the ceramic matrix composite turbine guide vane semi-finished product (the preparation process is the same as that of the silicon carbide ceramic matrix in the step (5)), and performing damage repair to obtain the ceramic matrix composite turbine guide vane.
Effect verification
The effect verification is carried out on the ceramic matrix composite turbine guide vane prepared in the embodiment 1-3, and the method specifically comprises the following steps: according to GJB 150.16A-2009 military equipment laboratory environmental test method part 16: vibration test ", the verification result is: under the conditions of a frequency band of 10-2000Hz and a total root mean square acceleration of 25grms, the tested blade has a complete structure, and abnormal phenomena such as cracking, layering, block dropping and the like do not occur, which shows that the method can improve the reliability of the component structure.
The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents, improvements and the like that fall within the spirit and principle of the present invention are intended to be included therein.
Claims (10)
1. The utility model provides a ceramic matrix composite turbine guide vane which characterized in that, turbine guide vane's material is ceramic matrix composite, and turbine guide vane's blade body and flange board be the integrated into one piece.
2. The ceramic matrix composite turbine vane of claim 1, wherein the reinforcement of the ceramic matrix composite is carbon fibers and/or silicon carbide fibers and the ceramic matrix is silicon carbide.
3. The method of making a ceramic matrix composite turbine vane of any one of claims 1-2, comprising the steps of:
(1) taking the molded surface of the inner cavity of the turbine guide blade as reference, preparing an inner mold (1) by using graphite, wherein the inner mold (1) is provided with a plurality of vent holes vertical to the molded surface;
(2) winding the woven fiber cloth on the outer surface of the inner mold (1), then sewing by using a sewing line made of the same material as the woven fiber cloth, wherein the wound woven fiber cloth after sewing is the prefabricated body (2) of the blade body, and then performing incision (3) treatment at positions 5-10mm away from two ends on the prefabricated body (2) of the blade body respectively, wherein the direction of the incision (3) is parallel to the longitudinal direction of the prefabricated body (2) of the blade body, and the length of the incision (3) is 10-50 mm;
(3) preparing 4 same plate-shaped prefabricated bodies (4) by adopting woven fiber cloth made of the same material as the prefabricated body (2) of the blade body, and arranging open holes (5) at corresponding positions of the blade body, wherein the shapes of the open holes (5) are the same as the cross sections of the prefabricated body (2) of the blade body;
(4) penetrating 4 sheets of plate-shaped prefabricated bodies (4) into the blade body prefabricated body (2) along the opening (5), wherein 2 sheets are positioned on two sides of the upper notch (3) of the blade body prefabricated body (2), the other 2 sheets are positioned on two sides of the lower notch (3) of the blade body prefabricated body (2), then respectively turning out the woven fiber cloth at the notch (3) of the blade body prefabricated body (2) along the outer surface of the blade body, clamping the turned-over edges (6) between the plate-shaped prefabricated bodies (4), respectively sewing the 2 sheets of plate-shaped prefabricated bodies (4) and the turned-over edges (6) clamped in the middle into a whole by using a sewing thread made of the same material as that of the blade body prefabricated body (2), and finally fixing the blade body prefabricated body (2) and the plate-shaped prefabricated body (4) by using an inner-shaped mold (1) to finish shaping;
(5) sequentially depositing an interface layer and a silicon carbide ceramic matrix on the surfaces of the shaped blade body prefabricated body (2) and the shaped plate-shaped prefabricated body (4), and removing the inner mould (1) to obtain a ceramic matrix composite turbine guide blade blank;
(6) and processing the ceramic matrix composite turbine guide vane blank to a design size, then continuously depositing silicon carbide on the surface of the ceramic matrix composite turbine guide vane blank, and performing damage repair to obtain the ceramic matrix composite turbine guide vane.
4. The method for preparing a ceramic matrix composite turbine vane of claim 3, wherein the diameter of the vent hole in step (1) is 2-5 mm.
5. The method of claim 3, wherein the woven fiber cloth material in step (2) is carbon fiber and/or silicon carbide fiber.
6. The method for preparing a ceramic matrix composite turbine guide vane according to claim 3, wherein in step (2), the distance between the notches (3) is 10-13mm.
7. The method of claim 3, wherein in step (5) the interfacial layer is a boron nitride interfacial layer.
8. The method for preparing a ceramic matrix composite turbine guide vane according to claim 3, wherein the silicon carbide ceramic matrix in the step (5) is prepared by the following steps: under the condition of pressure of 200-5000Pa, heating to 900-1200 ℃, keeping the temperature for 1-2h, introducing mixed gas of trichloromethylsilane, hydrogen and argon, depositing for 30-80h, keeping the temperature for 2h, and cooling to room temperature; this preparation process was performed 4-8 times.
9. The method of claim 3, wherein the machining in step (6) is by mechanical or laser machining.
10. The method of claim 3, wherein the silicon carbide deposition process of step (6) is the same as the silicon carbide ceramic substrate preparation process of step (5), and the process is performed 1-3 times.
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