CN110985134B - Aeroengine ceramic matrix composite fixed guider blade structure and forming method thereof - Google Patents

Aeroengine ceramic matrix composite fixed guider blade structure and forming method thereof Download PDF

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Publication number
CN110985134B
CN110985134B CN201911284946.4A CN201911284946A CN110985134B CN 110985134 B CN110985134 B CN 110985134B CN 201911284946 A CN201911284946 A CN 201911284946A CN 110985134 B CN110985134 B CN 110985134B
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rivet
blade body
ceramic matrix
limiting
upper edge
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CN110985134A (en
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涂建勇
王佳民
何江怡
陈旭
田录录
曹懿辰
成来飞
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Xian Xinyao Ceramic Composite Material Co Ltd
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Xi'an Golden Mountain Ceramic Composites Co ltd
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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
    • F01D9/00—Stators
    • 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
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • C04B35/515—Shaped ceramic products characterised by their composition; Ceramics compositions; Processing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on non-oxide ceramics
    • C04B35/56—Shaped ceramic products characterised by their composition; Ceramics compositions; Processing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on non-oxide ceramics based on carbides or oxycarbides
    • C04B35/565—Shaped ceramic products characterised by their composition; Ceramics compositions; Processing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on non-oxide ceramics based on carbides or oxycarbides based on silicon carbide
    • 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
    • C04B35/622—Forming processes; Processing powders of inorganic compounds preparatory to the manufacturing of ceramic products
    • 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
    • C04B35/622—Forming processes; Processing powders of inorganic compounds preparatory to the manufacturing of ceramic products
    • C04B35/626—Preparing or treating the powders individually or as batches ; preparing or treating macroscopic reinforcing agents for ceramic products, e.g. fibres; mechanical aspects section B
    • C04B35/628—Coating the powders or the macroscopic reinforcing agents
    • C04B35/62844—Coating fibres
    • C04B35/62857—Coating fibres with non-oxide ceramics
    • C04B35/62865—Nitrides
    • C04B35/62868—Boron nitride
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/005—Selecting particular materials
    • 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
    • C04B2235/00—Aspects relating to ceramic starting mixtures or sintered ceramic products
    • C04B2235/02—Composition of constituents of the starting material or of secondary phases of the final product
    • C04B2235/50—Constituents or additives of the starting mixture chosen for their shape or used because of their shape or their physical appearance
    • C04B2235/52—Constituents or additives characterised by their shapes
    • C04B2235/5208—Fibers
    • C04B2235/5216—Inorganic
    • C04B2235/524—Non-oxidic, e.g. borides, carbides, silicides or nitrides
    • C04B2235/5244—Silicon carbide
    • 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
    • C04B2235/00—Aspects relating to ceramic starting mixtures or sintered ceramic products
    • C04B2235/02—Composition of constituents of the starting material or of secondary phases of the final product
    • C04B2235/50—Constituents or additives of the starting mixture chosen for their shape or used because of their shape or their physical appearance
    • C04B2235/52—Constituents or additives characterised by their shapes
    • C04B2235/5208—Fibers
    • C04B2235/5252—Fibers having a specific pre-form
    • C04B2235/5256—Two-dimensional, e.g. woven structures
    • 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
    • C04B2235/00—Aspects relating to ceramic starting mixtures or sintered ceramic products
    • C04B2235/60—Aspects relating to the preparation, properties or mechanical treatment of green bodies or pre-forms
    • C04B2235/608—Green bodies or pre-forms with well-defined density
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material
    • F05D2230/31—Layer deposition
    • F05D2230/314—Layer deposition by chemical vapour deposition
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/603—Composites; e.g. fibre-reinforced
    • F05D2300/6033—Ceramic matrix composites [CMC]

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Ceramic Engineering (AREA)
  • Manufacturing & Machinery (AREA)
  • Materials Engineering (AREA)
  • Structural Engineering (AREA)
  • Organic Chemistry (AREA)
  • Inorganic Chemistry (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

The invention relates to a fixed guider blade structure and forming, in particular to a ceramic matrix composite fixed guider blade structure of an aero-engine and forming thereof, and belongs to the technical field of aero-engine fixed guider manufacturing. The defects that the existing CMC guider blade prefabricated body is unreasonable in design and poor in blade root strength are overcome, the turbine guider blade is prepared in an integrated assembly mode, parts are integrally assembled in a riveting mode, the assembly structure is welded by a SiC ceramic substrate, and the defect that the blade root strength is insufficient due to the fact that the prefabricated body is conventionally sewn and formed is overcome. The spare of many parts, the characteristics of preferred assembly can be realized, the condemned risk of whole component that can avoid single part to scrap causes reduces the preparation risk, and reduce cost improves the component quality.

Description

Aeroengine ceramic matrix composite fixed guider blade structure and forming method thereof
Technical Field
The invention relates to a fixed guider blade structure and forming, in particular to an assembly structure and a forming method of a ceramic matrix composite fixed guider blade for an aero-engine, and belongs to the technical field of aero-engine fixed guider preparation.
Background
The military and civil aircraft has increasingly urgent need for high-performance aircraft engines, and there are two main ways to improve the performance of the engines, one is to improve the pressure ratio of a gas compressor, and the other is to improve the temperature of gas at the inlet of a turbine. As turbine inlet gas temperatures increase, the high temperature components of the engine are subjected to greater heat loads. The gas temperature before the turbine of the engine with the active thrust-weight ratio of 10 reaches 1850-1950K, and can generate one time of more thrust than the previous generation of aeroengines; the thrust-weight ratio of the fifth generation aero-engine in the future can be as high as 15-20, the gas temperature before the turbine can be as high as 2200-2400K, and the temperature is far beyond the heat resistance limit of the current turbine and turbine front-end component material.
Compared with high-temperature alloy, the Ceramic Matrix Composite (CMC) can bear higher temperature and can obviously reduce cooling airflow; the strength under the high-temperature working condition is high, the modulus is high, the damping performance is good, the impact resistance is excellent, and the service life is long; meanwhile, the density of the CMC material is 2.0-2.5 g/cm3The high-temperature alloy is 1/4-1/3, the structure mass can be greatly reduced, and the thrust-weight ratio of the engine can be improved by simple replacement; therefore, the CMC material becomes the most potential substitute and upgrading material for advanced engine hot-end components, and has great potential application on engine hot-end static and rotor components.
The turbine guider of the aircraft engine is positioned at the front end of the turbine rotor, and the temperature environment is relatively severer. The document "Halbig M, Jaskowiak M, Kiser J, et al, Evaluation of ceramic matrix composite technology for air turbine engine applications 51st AIAA air space science Evaluation including the New thermal area form and air space exposure, 2013[ C ]" verifies the manufacturability of complex parts including high pressure turbine blades and evaluates their performance and durability under simulated engine operating conditions.
The document "Takashi A, Takeshi N, Kooun T, et al, Research of CMC Application to Turbine Components [ J ]. IHI Engineering Review, 2005,38(2): 58-62" discloses a report on CMC low pressure Turbine nozzle vanes developed by IHI corporation of Japan. The blade preform is formed by partially sewing an upper edge plate preform, a blade body preform and a lower edge plate preform 3, sewing fibers are concentrated at the position of a blade root formed by vertically intersecting the upper edge plate and the blade body, the blade root is the position with the most concentrated working condition stress in a service state, however, the included angle between the blade body and the upper edge plate and the included angle between the blade body and the lower edge plate are about 90 degrees, the sewing fibers can be cut off in a large quantity in the subsequent blade forming process, the fiber continuity is damaged, the structural strength of the blade root of the guider is weakened, the performance of the CMC fixed guider is restricted, and the service strength and the reliability of the CMC blade are influenced.
Disclosure of Invention
In order to overcome the structural defects that the existing CMC (ceramic matrix composite) guider blade preform is unreasonable in design and poor in blade root strength, the invention provides a novel ceramic matrix composite fixed guider blade structure and a forming method thereof.
The invention provides a ceramic matrix composite fixed guider blade structure of an aero-engine, which comprises a blade body, an upper edge plate component fixed at the top of the blade body and a lower edge plate fixed at the bottom of the blade body, and is characterized in that: the riveting device also comprises a limit rivet and a riveting rivet;
the upper edge plate component comprises a locking flat plate and an upper edge plate main body;
the blade body comprises a blade body and a boss arranged on the upper end face of the blade body, and the boss and the blade body are integrally arranged; a limiting hole penetrating through the boss is formed along the side wall of the boss;
the upper edge plate main body is provided with a blade body limiting notch; the shape of the blade body limiting notch is matched with that of the boss;
the lug boss penetrates through the notch; the limiting rivet penetrates through the limiting hole and is lapped on the upper surface of the upper edge plate main body;
the locking flat plate covers the upper surface of the upper flange plate main body and is fixed on the upper surface of the upper flange plate main body through riveting rivets; the locking flat plate is used for limiting the freedom degree of the limit rivet;
the blade body, the upper edge plate main body, the limiting rivet, the locking flat plate and the riveting rivet are all made of ceramic matrix composite materials; and the whole of the guider blade structure is deposited with a SiC ceramic matrix; depositing a SiC ceramic matrix at the contact position of the limiting rivet, the boss and the upper edge plate main body; and depositing a SiC ceramic matrix at the contact position of the riveting rivet and the locking flat plate. After the assembly is finished, the components are finished in a chemical vapor deposition furnaceThe purpose of bulk deposition of silicon carbide substrates is:
Figure DEST_PATH_IMAGE001
densifying the monolith;
Figure 863751DEST_PATH_IMAGE002
eliminating assembly gaps during assembly;
Figure DEST_PATH_IMAGE003
and (4) loosening the rivet (4) and the boss (7).
Furthermore, the blade body comprises a first curved blade body, a second curved blade body and a third curved blade body, wherein the first curved blade body and the second curved blade body are in transition connection through the third curved blade body and are integrally arranged;
the two bosses are respectively arranged on the upper end surfaces of the first curved blade body and the second curved blade body.
Furthermore, the two limit rivets respectively penetrate through the limit holes on the bosses at the upper end faces of the first curved blade body and the second curved blade body and are lapped on the upper surface of the upper edge plate main body.
Furthermore, a rivet limiting notch is formed in the locking flat plate, and the inner edge of the rivet limiting notch is in contact fit with the limiting rivet to limit the horizontal degree of freedom of the limiting rivet.
Further, the ceramic matrix composite is stacked by a plurality of layers of plain cloth; the blade body, the upper edge plate main body, the limiting rivet, the locking flat plate and the riveting rivet are all formed by cutting along the stacking direction of the plain cloth.
The invention also provides a forming method of the ceramic matrix composite fixed guider blade structure of the aero-engine, which comprises the following steps:
preparing a blade body, an upper edge plate main body, a limiting rivet, a locking flat plate and a riveting rivet by using a ceramic matrix composite;
secondly, inserting a boss on the blade body into a blade body limiting notch on the upper edge plate main body, and then inserting a limiting rivet into a limiting hole to realize the positioning of the upper edge plate main body;
thirdly, placing the product positioned and molded in the second step into a chemical vapor deposition furnace (CVI), and depositing a SiC ceramic substrate on the product by adopting a CVI process to complete the fixation of the limit rivet;
covering the locking flat plate on the upper surface of the upper edge plate main body, and locking by riveting rivets; the inner edge of a rivet limit notch of the locking flat plate is in contact fit with a limit rivet, and the horizontal degree of freedom of the limit rivet is limited;
and step five, placing the product formed in the step four in a chemical vapor deposition furnace, and depositing the SiC ceramic matrix on the whole product by adopting a CVI (chemical vapor deposition) process.
Further, the first step is specifically as follows:
step 1.1, preparing a fiber preform:
weaving 2D plain cloth by adopting SiC fibers, cutting the plain cloth into proper size according to the size of a part, stacking a plurality of layers of plain cloth, puncturing the SiC fibers in the stacking direction of the plain cloth to form a fiber preform, and fixing and molding the fiber preform by utilizing a mold;
step 1.2, preparing an interface layer:
placing the fiber preform fixedly formed in the step 1.1 in a chemical vapor deposition furnace, and preparing an interface layer on the fiber surface of the preform;
step 1.3, preparing a SiC ceramic matrix:
placing the product prepared in the step 1.2 in a chemical vapor deposition furnace CVI, and depositing a SiC ceramic matrix on the product by adopting a CVI process;
step 1.4, part processing:
and (4) placing the product prepared in the step (1.3) on processing equipment, and cutting along the plain weave cloth stacking direction to process the blade body, the upper edge plate main body, the limiting rivet, the locking flat plate and the riveting rivet.
Further, the process conditions for preparing the interface layer in step 1.2 are as follows:
the deposition temperature of the boron nitride BN interface layer is 400-1200 ℃, the deposition furnace is vacuumized to 3-50 kPa, 60-100L/min of Ar gas is used as protective gas, the flow of precursor gas of the interface layer is 10-500L/min, and the deposition time is 20-50 h;
step 1.3, the preparation of the SiC ceramic matrix and the deposition of the SiC ceramic matrix in the step three have the following process conditions:
the temperature is 1200-1500 ℃, the deposition furnace is vacuumized to 3-50 kPa, and the H is 60-100L/min2The gas is used as carrier gas, the gas flow of the ceramic matrix precursor is 10-500L/min, the single deposition time is 100-150 h, and the densification deposition is circulated for multiple times until the density is more than or equal to 2.0g/cm3。
Further, in the step 1.4, the processing equipment adopts a common multi-axis numerical control machine tool, and the processing cutter adopts cubic boron nitride or diamond.
The invention has the beneficial effects that:
1. according to the invention, the turbine guider blade is prepared in an integrated assembly mode, parts are integrally assembled in a riveting mode, and an assembly structure adopts a SiC ceramic substrate to realize welding, so that the defect of insufficient strength of the blade root caused by conventional sewing and forming of a prefabricated body is avoided;
2. the invention has the characteristics of realizing backup of multiple parts and optimal assembly, can avoid the risk of scrapping the whole component caused by scrapping a single part, reduces the preparation risk, reduces the cost and improves the component quality.
3. The scheme provided by the invention has strong technological adaptability and can be prepared in batch and in industrialization.
4. The invention makes full use of the available space between the upper edge plate component and the casing, so that the complex SiC/SiC high-pressure guide blade has good manufacturability, the reliable connection between the edge plate and the blade is realized, and the structure can be fully applied to small and medium-sized engines, such as turboshaft engines, turbojet engines and turbofan engines.
Drawings
FIG. 1 is a schematic view of the blade body structure of the present invention;
FIG. 2 is a schematic view of the relative position of the upper edge plate and the blade body of the present invention;
FIG. 3 is a schematic view of the attachment of the upper platform to the blade body according to the present invention;
FIG. 4 is a schematic view of the upper flange riveting and locking plate of the present invention;
FIG. 5 is a schematic view of the assembly of the upper platform and the blade body of the present invention;
in the figure, 1-blade body, 11-first curved blade body, 12-second curved blade body, 13-third curved blade body, 2-limiting hole, 3-upper edge plate main body, 31-blade body limiting notch, 32-rivet limiting notch, 4-limiting rivet, 5-locking flat plate, 6-riveting rivet, 7-boss and 8-upper edge plate component.
Detailed Description
The present invention will be described in detail below with reference to the accompanying drawings and specific embodiments. The detailed description of the present invention is provided to further explain the concept of the present invention, the technical problems to be solved, and the features and effects of the technical solutions. The description of the embodiments is not intended to limit the present invention. Further, the technical features according to the embodiments of the present invention may be combined with each other as long as they do not conflict with each other.
The invention relates to a ceramic matrix composite fixed guider blade structure of an aero-engine, which comprises a blade body 1, an upper edge plate assembly 8, a locking flat plate 5, a limiting rivet 4 and a riveting rivet 6.
The blade body 1 comprises a blade body and a boss 7 arranged on the upper end face of the blade body, and the boss 7 and the blade body are integrally arranged; and a limiting hole 2 penetrating through the boss is formed along the side wall of the boss 7. As shown in fig. 1, the blade body of the present invention includes a first curved blade body 11, a second curved blade body 12, and a third curved blade body 13, the first curved blade body 11 and the second curved blade body 12 are transitionally connected by the third curved blade body 13, the cross-section forms a V-shaped structure, and the first curved blade body 11, the second curved blade body 12, and the third curved blade body 13 are integrally disposed. The invention comprises two bosses 7 which are respectively arranged on the upper end surfaces of a first curved blade body 11 and a second curved blade body 12.
Referring to fig. 2 to 5, the upper edge plate assembly 8 of the present invention includes an upper edge plate main body 3 and a locking plate 5; the middle part of the upper edge plate main body 3 is provided with a blade body limiting notch 31, and the shape of the blade body limiting notch 31 is the same as the cross section of the upper end of the boss 7; the locking plate 5 is provided with a rivet limiting notch 32. During assembly, the boss 7 is inserted into the corresponding blade body limiting notch 31, and the limiting rivet 4 passes through the limiting hole 2 and is lapped on the upper surface of the upper edge plate main body 3 to limit the relative positions of the blade body 1 and the upper edge plate main body 3. Then, performing chemical vapor deposition (CVI) or preparing a SiC ceramic matrix by adopting other processes, and performing 'welding' treatment on the riveting structure of the limit pin 4 to deposit the SiC ceramic matrix at the reserved connecting gap between the limit rivet 4 and the boss 7 as well as the reserved connecting gap between the limit rivet 4 and the upper edge plate main body 3; then, two locking flat plates 5 are covered and riveted on the upper surface of the upper flange plate body 3 by riveting rivets 6, and the horizontal freedom degree of the limit rivet 4 is limited. Then preparing SiC ceramic matrix by adopting Chemical Vapor Infiltration (CVI) process or other processes, depositing the SiC ceramic matrix at the joint of the riveting rivet 6, and carrying out 'welding' treatment on the riveting structure of the riveting rivet 6; meanwhile, the assembly clearance during assembly is eliminated, and the integral component is densified again to complete assembly. After the blade body 1, the upper edge plate main body 3, the locking flat plate 5, the limiting rivet 4 and the riveting rivet 6 are assembled step by step, deposition is carried out for many times (80 hours/time, about 6 times), and excess is mechanically ground for many times through the middle. Carrying out nondestructive inspection on subsequent products through X-ray or infrared thermal wave imaging: inclusions, delamination, holes, cracks, density uniformity, etc., if found, were further detected and analyzed by CT.
Example one
The CMC fixed guide vane size of the present embodiment: 60mm long, 50mm wide and 120mm high. In this embodiment, the ceramic matrix composite material is made of SiC fiber, and the raw materials used in the preparation process are trichloromethylsilane and H2Ar gas, etc.
The method comprises the following specific steps:
(1) preparing a prefabricated body: the SiC fiber is woven into 2D plain cloth, and other prefabricated body types such as 2.5D, 3D and the like can also be adopted. Cutting the plain cloth into proper size according to the size of the part, stacking the multilayer plain cloth, puncturing in the stacking direction of the plain cloth, and forming the SiC fiber preform by using the same SiC fiber as the puncturing fiber. And fixing and molding the SiC fiber preform by using a mold.
(2) Preparing an interface layer: and (3) placing the prefabricated body in the step (1) in a chemical vapor deposition furnace, and preparing an interface layer. The deposition temperature of the boron nitride BN interface layer is 400-1200 ℃, the deposition furnace is vacuumized to 3-50 kPa, 60-100L/min of Ar gas is used as protective gas, the flow of precursor gas of the interface layer is 10-500L/min, and the deposition time is 20-50 h.
(3) Preparing a SiC ceramic matrix: placing the product prepared in the step (2) in a chemical vapor deposition furnace (CVI), and depositing a SiC ceramic matrix on the product by adopting a CVI process;
the deposition temperature of the SiC ceramic matrix is 1200-1500 ℃, and the deposition furnace is vacuumized to H of 3-50 kPa, 60-100L/min2Gas is used as carrier gas, the flow rate of the ceramic matrix precursor gas is 10-500L/min, the single deposition time is 100-150 h, the densification deposition is carried out for multiple cycles, and the density is more than or equal to 2.0g/cm3After that, the next process is performed.
(4) Processing parts: and (4) placing the product obtained in the step (3) on a processing device, wherein the processing device adopts a common multi-shaft numerical control machine tool, and the processing cutter adopts special cutters such as cubic boron nitride, diamond and the like. Processing the outline of the blade body 1 and the limiting hole 2; processing a blade body limiting notch 31 and a profile on the upper edge plate main body 3; and processing accessories such as a limit rivet 4 and the like. And cutting along the stacking direction of the plain cloth during processing.
(5) Repairing processing damage: after mechanical processing, placing the product processed in the step (4) in a chemical vapor deposition furnace (CVI) by using the process in the step (3), and depositing a SiC ceramic matrix on the product by adopting the CVI process; although the processing area can cause tiny damage to products or local fiber breakage, the porosity of the composite material can be increased at the processing position, and a high-density area can be generated in the processing area through subsequent deposition, so that the repairing effect on the whole structure is achieved.
(6) The upper edge plate main body is assembled with the blade body. Inserting a boss 7 at the upper end of the blade body 1 into a blade body limiting notch 31 of the upper edge plate main body 3, and inserting a limiting pin 4 into a limiting hole 2; placing the assembled product in a chemical vapor deposition furnace (CVI), depositing a SiC ceramic substrate on the product by the CVI process by utilizing the process in the step (3), and performing 'welding' treatment on the riveting structure of the limit pin 4; the locking plate 5 is then riveted to the upper flange body 3, again by means of the process using step (3), the riveted structure being "welded".

Claims (4)

1. The utility model provides an aeroengine ceramic matrix composite fixes director blade structure, includes blade (1), fixes last reason flange components (8) at blade (1) top and fixes the lower flange board in blade (1) bottom, its characterized in that: the riveting device also comprises a limit rivet (4) and a riveting rivet (6);
the upper edge plate component (8) comprises a locking flat plate (5) and an upper edge plate main body (3);
the blade body (1) comprises a blade body and a boss (7) arranged on the upper end face of the blade body, and the boss (7) and the blade body are integrally arranged; a limiting hole (2) penetrating through the boss is formed along the side wall of the boss (7); the blade body comprises a first curved blade body (11), a second curved blade body (12) and a third curved blade body (13), wherein the first curved blade body (11) and the second curved blade body (12) are in transitional connection through the third curved blade body (13) and are integrally arranged; the two bosses are respectively arranged on the upper end faces of the first curved blade body (11) and the second curved blade body (12);
a blade body limiting notch (31) is formed in the upper edge plate main body (3); the shape of the blade body limiting notch (31) is matched with that of the boss (7);
the lug boss (7) penetrates through the notch; the limiting rivets (4) respectively penetrate through limiting holes (2) on bosses at the upper end faces of the first curved blade body (11) and the second curved blade body (12) and are lapped on the upper surface of the upper edge plate main body (3);
the locking flat plate (5) covers the upper surface of the upper edge plate main body (3), and the locking flat plate (5) is fixed on the upper surface of the upper edge plate main body (3) through a riveting rivet (6); the locking flat plate (5) is used for limiting the degree of freedom of the limit rivet (4); a rivet limiting notch (32) is formed in the locking flat plate (5), the inner edge of the rivet limiting notch (32) is in contact fit with the limiting rivet (4), and the horizontal degree of freedom of the limiting rivet (4) is limited;
the blade body (1), the upper edge plate main body (3), the limiting rivet (4), the locking flat plate (5) and the riveting rivet (6) are all made of ceramic matrix composite materials; the ceramic matrix composite is stacked by a plurality of layers of plain cloth; the leaf body (1), the upper edge plate main body (3), the limiting rivet (4), the locking flat plate (5) and the riveting rivet (6) are formed by cutting along the plain cloth stacking direction;
integrally depositing a SiC ceramic matrix on the guide vane structure;
depositing SiC ceramic matrix at the reserved connection gap between the limit rivet (4) and the boss (7) and the upper edge plate main body (3); and SiC ceramic matrix is deposited at the reserved connection gap between the riveting rivet (6) and the locking flat plate (5).
2. The aero engine ceramic matrix composite fixed guide vane structure of claim 1 wherein: the number of the limiting rivets (4) is two.
3. A method of forming an aero engine ceramic matrix composite fixed guide vane structure as defined in claim 1 comprising the steps of:
preparing a blade body (1), an upper edge plate main body (3), a limiting rivet (4), a locking flat plate (5) and a riveting rivet (6) by using a ceramic matrix composite material;
secondly, inserting a boss (7) on the blade body (1) into a blade body limiting notch (31) on the upper edge plate main body (3), and then inserting a limiting rivet (4) into a limiting hole (2) to realize the positioning of the upper edge plate main body (3);
thirdly, placing the product positioned and molded in the second step into a chemical vapor deposition furnace (CVI), and depositing a SiC ceramic substrate on the product by adopting a CVI process to complete the fixation of the limit rivet (4);
covering the upper surface of the upper edge plate main body (3) with a locking flat plate (5), and locking by riveting rivets (6); the inner edge of a rivet limit notch (32) of the locking flat plate (5) is in contact fit with the limit rivet (4) to limit the horizontal degree of freedom of the limit rivet (4);
placing the product formed in the fourth step into a chemical vapor deposition furnace, and depositing a SiC ceramic matrix on the whole product by adopting a CVI (chemical vapor deposition) process;
the first step is specifically as follows:
step 1.1, preparing a fiber preform:
weaving 2D plain cloth by adopting SiC fibers, cutting the plain cloth according to the size of a part, stacking a plurality of layers of plain cloth, puncturing the SiC fibers in the stacking direction of the plain cloth to form a fiber preform, and fixing and molding the fiber preform by utilizing a mold;
step 1.2, preparing an interface layer:
placing the fiber preform fixedly formed in the step 1.1 in a chemical vapor deposition furnace for preparing an interface layer;
step 1.3, preparing a SiC ceramic matrix:
placing the product prepared in the step 1.2 in a chemical vapor deposition furnace CVI, and depositing a SiC ceramic matrix on the product by adopting a CVI process;
step 1.4, part processing:
placing the product prepared in the step 1.3 on processing equipment, and cutting along the plain cloth stacking direction to process the blade body (1), the upper edge plate main body (3), the limiting rivet (4), the locking flat plate (5) and the riveting rivet (6);
step 1.2 the process conditions for preparing the interface layer are as follows:
the deposition temperature of the boron nitride BN interface layer is 400-1200 ℃, the deposition furnace is vacuumized to 3-50 kPa, 60-100L/min of Ar gas is used as protective gas, the flow of precursor gas of the interface layer is 10-500L/min, and the deposition time is 20-50 h;
step 1.3, the process conditions of the preparation of the SiC ceramic matrix and the deposition of the SiC ceramic matrix in the step three are as follows:
the temperature is 1200-1500 ℃, the deposition furnace is vacuumized to 3-50 kPa, and the H is 60-100L/min2Gas is used as carrier gas, the flow rate of the ceramic matrix precursor gas is 10-500L/min, the single deposition time is 100-150 h, and the densification deposition is circulated for multiple times until the density is more than or equal to 2.0g/cm3。
4. The method of forming an aircraft engine ceramic matrix composite fixed guide vane structure of claim 3, wherein: and (1.4) adopting a common multi-axis numerical control machine tool as processing equipment, and adopting cubic boron nitride or diamond as a processing cutter.
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