CN111102017B - 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 PDFInfo
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- CN111102017B CN111102017B CN201911283617.8A CN201911283617A CN111102017B CN 111102017 B CN111102017 B CN 111102017B CN 201911283617 A CN201911283617 A CN 201911283617A CN 111102017 B CN111102017 B CN 111102017B
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- 239000011153 ceramic matrix composite Substances 0.000 title claims abstract description 31
- 238000000034 method Methods 0.000 title claims abstract description 30
- 239000000919 ceramic Substances 0.000 claims abstract description 34
- 239000011159 matrix material Substances 0.000 claims abstract description 29
- 238000002360 preparation method Methods 0.000 claims abstract description 9
- 239000000758 substrate Substances 0.000 claims abstract description 4
- 238000000151 deposition Methods 0.000 claims description 37
- 239000000835 fiber Substances 0.000 claims description 26
- 230000008021 deposition Effects 0.000 claims description 23
- 239000007789 gas Substances 0.000 claims description 21
- 238000012545 processing Methods 0.000 claims description 21
- 238000005229 chemical vapour deposition Methods 0.000 claims description 20
- 239000004744 fabric Substances 0.000 claims description 19
- 238000005520 cutting process Methods 0.000 claims description 8
- 239000000463 material Substances 0.000 claims description 8
- 229910052582 BN Inorganic materials 0.000 claims description 6
- PZNSFCLAULLKQX-UHFFFAOYSA-N Boron nitride Chemical compound N#B PZNSFCLAULLKQX-UHFFFAOYSA-N 0.000 claims description 6
- 239000002243 precursor Substances 0.000 claims description 6
- 239000012159 carrier gas Substances 0.000 claims description 3
- 238000000280 densification Methods 0.000 claims description 3
- 229910003460 diamond Inorganic materials 0.000 claims description 3
- 239000010432 diamond Substances 0.000 claims description 3
- 238000000465 moulding Methods 0.000 claims description 3
- 230000000149 penetrating effect Effects 0.000 claims description 3
- 230000001681 protective effect Effects 0.000 claims description 3
- 230000007704 transition Effects 0.000 claims description 2
- 238000009941 weaving Methods 0.000 claims description 2
- 238000009958 sewing Methods 0.000 abstract description 5
- 238000003466 welding Methods 0.000 abstract description 5
- 230000007547 defect Effects 0.000 abstract description 2
- 230000008595 infiltration Effects 0.000 abstract description 2
- 238000001764 infiltration Methods 0.000 abstract description 2
- 239000000126 substance Substances 0.000 abstract description 2
- 239000000047 product Substances 0.000 description 17
- 238000005516 engineering process Methods 0.000 description 3
- 239000000956 alloy Substances 0.000 description 2
- 229910045601 alloy Inorganic materials 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 238000001816 cooling Methods 0.000 description 1
- 238000013016 damping Methods 0.000 description 1
- 230000032798 delamination Effects 0.000 description 1
- 238000013461 design Methods 0.000 description 1
- 238000011156 evaluation Methods 0.000 description 1
- 239000012467 final product Substances 0.000 description 1
- 238000003384 imaging method Methods 0.000 description 1
- 238000007689 inspection Methods 0.000 description 1
- 238000003754 machining Methods 0.000 description 1
- 239000002994 raw material Substances 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 238000012552 review Methods 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
- 230000007847 structural defect Effects 0.000 description 1
- DWAWYEUJUWLESO-UHFFFAOYSA-N trichloromethylsilane Chemical compound [SiH3]C(Cl)(Cl)Cl DWAWYEUJUWLESO-UHFFFAOYSA-N 0.000 description 1
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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
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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
- C04B35/515—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
-
- 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
-
- 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
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Ceramic Engineering (AREA)
- Materials Engineering (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Manufacturing & Machinery (AREA)
- Structural Engineering (AREA)
- Organic Chemistry (AREA)
- Inorganic Chemistry (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention relates to a fixed guider blade structure and forming, in particular to an aero-engine ceramic matrix composite fixed guider blade structure and forming thereof, and belongs to the technical field of aero-engine fixed guider preparation. The turbine guide vane is prepared by adopting an integrated assembly mode, parts are integrally assembled by adopting a riveting mode, and an assembly structure adopts a SiC ceramic substrate to realize welding, so that the defect of insufficient strength of a vane root caused by the conventional sewing forming of a prefabricated body is avoided; according to the invention, after the bulges on the first lower edge plate and the second lower edge plate are directly subjected to limit riveting with the blade, the first lower edge plate and the second lower edge plate are riveted with the third lower edge plate and the fourth lower edge plate through riveting rivets, and a SiC ceramic matrix is deposited on the whole product by adopting a CVI (chemical vapor infiltration) process, so that the integrated preparation is completed. The SiC/SiC high-pressure guide blade has good manufacturability, the edge plate and the blade are simpler to prepare and more reliable to install, and the structure can be fully applied to small engines, such as civil turboshaft engines and turbojet engines.
Description
Technical Field
The invention relates to a fixed guider blade structure and forming, in particular to a ceramic matrix composite fixed guider blade assembly structure and a forming method 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 thermal loads. The gas temperature before the turbine of the engine with the active thrust-weight ratio of 10 in the developed country 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 weight-pushing ratio of the engine can be improved by simply replacing 1/4-1/3 of high-temperature alloy, so that the structural mass can be greatly reduced; therefore, the CMC material becomes the most potential substitute and upgrading material for the hot-end part of the advanced engine, and has great application potential on the static and rotor parts at the hot end of the engine.
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 AIAAAAAero space Sciences machining including the New Horizons form and Aero 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 literature "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, 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 of unreasonable design and poor strength of a blade root of the conventional CMC guider blade preform, 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 fixed at the top of the blade body and a lower edge plate component fixed at the bottom of the blade body, and is characterized in that: also comprises a riveting rivet;
the blade body comprises a blade body and a limiting table arranged on the lower end face of the blade body, and the limiting table and the blade body are integrally arranged; 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 limiting tables are respectively arranged on the lower end surfaces of the first curved blade body and the second curved blade body; a limiting hole penetrating through the limiting table is formed along the side wall of the limiting table;
the lower edge plate assembly comprises a first lower edge plate, a second lower edge plate, a third lower edge plate and a fourth lower edge plate; the first lower edge plate and the second lower edge plate are both provided with bulges which can be inserted into the limiting holes;
the bulges on the first lower edge plate and the second lower edge plate are respectively inserted into the two limiting holes, and the two limiting tables penetrate through the gap formed by the first lower edge plate and the second lower edge plate;
the third lower edge plate and the fourth lower edge plate respectively cover the outer surfaces of the first lower edge plate and the second lower edge plate, and the third lower edge plate and the fourth lower edge plate are respectively fixed on the outer surfaces of the first lower edge plate and the second lower edge plate through riveting rivets; the third lower edge plate is provided with an engine connecting hole;
the blade body, the lower edge plate assembly and the riveting rivet are all made of ceramic matrix composite materials;
integrally depositing a SiC ceramic matrix on the guide vane structure; depositing a SiC ceramic matrix at the reserved connection gap between the protrusion and the limiting hole; and SiC ceramic matrix is deposited at the reserved connecting gaps of the riveting rivet, the third lower edge plate and the fourth lower edge plate.
Furthermore, a notch is formed in the third lower edge plate, and the inner edge of the notch is in contact fit with the outer edges of the two limiting tables.
Further, the ceramic matrix composite is stacked by a plurality of layers of plain cloth; the blade body, the lower edge plate assembly and the riveting rivet are all formed by cutting along the stacking direction of the fiber 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, a lower edge plate assembly and a riveting rivet by using a ceramic matrix composite;
secondly, inserting the bulges on the first lower edge plate and the second lower edge plate into the two limiting holes respectively, and enabling the two limiting tables to penetrate through the gaps formed by the first lower edge plate and the second lower edge plate; positioning the first lower edge plate and the second lower edge plate;
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 bulge;
step four, covering the surfaces of the first lower edge plate and the second lower edge plate with a third lower edge plate and a fourth lower edge plate respectively, and locking by riveting rivets; the inner edges of the notches are in contact fit with the outer edges of the two limiting tables;
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 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:
and (4) placing the product prepared in the step 1.3 on processing equipment, and cutting along the stacking direction of the fiber cloth to process the blade body, the lower edge plate assembly 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/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。
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. for the small-sized aeroengine with the small available space between the lower edge plate and the casing and the limited size of the blade capable of protruding out of the lower edge plate, the invention adopts the technical scheme that after the protrusions on the first lower edge plate and the second lower edge plate are directly in limit riveting with the blade, the protrusions are riveted with the third lower edge plate and the fourth lower edge plate through riveting rivets, and the SiC ceramic matrix is deposited on the whole product by adopting the CVI process, thereby completing the integrated preparation. The invention relates to a preparation method of a SiC/SiC high-pressure guide blade with a complex structure, which is a great problem in the world at present. The SiC/SiC high-pressure guide blade has good manufacturability, the edge plate and the blade are simpler to prepare, more reliable to mount, higher in structural strength, and easier to guarantee the dimensional accuracy of the final product. The structure can be fully applied to small engines, such as civil turboshaft engines and turbojet engines.
3. The invention has the characteristics of realizing multi-part backup 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.
4. The scheme provided by the invention has strong technological adaptability and can be prepared in batch and in industrialization.
Drawings
FIG. 1 is a schematic view of the blade structure of the present invention;
FIG. 2 is a schematic view of the relative position of the lower edge plate and the blade body of the present invention;
FIG. 3 is a schematic view of the attachment of the lower platform to the blade body according to the present invention;
FIG. 4 is a schematic view of the lower flange securing and locking plate of the present invention;
FIG. 5a is a first view of the assembly of the locking plate and the lower flange of the present invention;
FIG. 5b is a second view of the locking plate and lower flange assembly of the present invention;
FIG. 6 is a schematic view of the lower flange, locking plate and blade body assembly of the present invention;
in the figure, 01-blade body, 11-first curved blade body, 12-second curved blade body, 13-third curved blade body, 02-lower edge plate component, 021-first lower edge plate, 022-second lower edge plate, 023-third lower edge plate, 024-fourth lower edge plate, 03-engine connecting hole, 031-notch, 04-riveting rivet, 05-limiting platform, 06-limiting hole and 07-bulge.
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 blade structure of a ceramic matrix composite fixed guider of an aero-engine, which comprises a blade body 01, an upper edge plate, a lower edge plate assembly 02 and riveting rivets 04.
The blade body 01 comprises a blade body and a limiting table 05 arranged on the lower end face of the blade body, and the limiting table 05 and the blade body are integrally arranged; and a limiting hole 06 penetrating through the limiting table is formed along the side wall of the limiting table 05. 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 limiting tables 05 which are respectively arranged on the lower end surfaces of a first curved blade body 11 and a second curved blade body 12.
2-6, lower edge plate assembly 02 of the present invention includes a first lower edge plate 021, a second lower edge plate 022, a third lower edge plate 023, and a fourth lower edge plate 024; the first lower edge plate 021 and the second lower edge plate 022 are respectively provided with a bulge 07 which can be inserted into the limiting hole 06; the protrusions on the first lower edge plate 021 and the second lower edge plate 022 are inserted into the two limiting holes 06, respectively, and the two limiting platforms 05 pass through the notch formed by the first lower edge plate 021 and the second lower edge plate 022. The third lower flange 023 is an L-shaped plate, a horizontal plane of the third lower flange 023 is provided with a notch 031, and a vertical plane of the third lower flange 023 is provided with an engine connector 03.
During assembly, the protrusions 07 on the first lower edge plate 021 and the second lower edge plate 022 are respectively inserted into the two corresponding limiting holes 06; the limiting table 05 is inserted into the corresponding notch; the relative positions of the blade body 01 and the first lower edge plate 021 and the second lower edge plate 022 are restricted. Then, performing chemical vapor deposition (CVI) technology or preparing the SiC ceramic matrix by adopting other technologies, and performing 'welding' treatment on the riveting structure of the bulge 07 to deposit the SiC ceramic matrix at the reserved connecting gap between the bulge 07 and the limiting table 05 as well as the limiting hole 06; then, the third and fourth lower flanges 023 and 024 are respectively covered and riveted to the surfaces of the first and second lower flanges 021 and 022 by riveting rivets 04. Then preparing the SiC ceramic matrix by adopting a Chemical Vapor Infiltration (CVI) process or other processes, depositing the SiC ceramic matrix at the joint of the riveting rivet 04, and carrying out 'welding' treatment on the riveting structure of the riveting rivet 04; meanwhile, the assembly clearance during assembly is eliminated, and the integral component is densified again to complete assembly. After the blade body 01, the lower edge plate assembly 02 and the riveting rivet 04 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 sizes according to the sizes of parts, stacking the multilayer plain cloth, puncturing in the stacking direction of the plain cloth, wherein the puncturing fibers adopt the same SiC fibers to form a SiC fiber preform. 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/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, the densification deposition is circulated for multiple times, 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 01 and a limiting hole 06; processing a blade body limiting hole 06 and a profile on the lower edge plate component 02; and processing accessories such as riveting rivets 04 and the like. And during processing, cutting along the stacking direction of the plain cloth.
(5) Repairing processing damage: after mechanical processing, placing the product processed in the step (4) in a chemical vapor deposition furnace (CVI) by utilizing 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 lower flange plate is assembled with the blade body. Inserting the protrusion 07 into the limiting hole 06, and inserting the limiting platform 05 at the lower end of the blade body 01 into a gap formed between the first lower edge plate 021 and the second lower edge plate 022; placing the assembled product in a chemical vapor deposition furnace (CVI), depositing a SiC ceramic matrix on the product by the process of the step (3) and adopting the CVI process, and carrying out 'welding' treatment on the riveted structure of the bulge 07; the third and fourth lower edge plates 023 and 024 are then riveted to the first and second lower edge plates 021 and 022, respectively, and the riveted structure is "welded" also by the process using step (3).
Claims (5)
1. The utility model provides an aeroengine ceramic matrix composite fixes director blade structure, includes blade (01), fixes the last flange plate at blade (01) top and fixes lower flange plate subassembly (02) in blade (01) bottom, its characterized in that: also comprises a riveting rivet (04);
the blade body (01) comprises a blade body and a limiting table (05) arranged on the lower end face of the blade body, and the limiting table (05) and the blade body are integrally arranged; 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 transition connection through the third curved blade body (13) and are integrally arranged; the two limiting tables (05) are respectively arranged on the lower end faces of the first curved blade body (11) and the second curved blade body (12); a limiting hole (06) penetrating through the limiting table is formed along the side wall of the limiting table (05);
the lower edge plate assembly (02) comprises a first lower edge plate (021), a second lower edge plate (022), a third lower edge plate (023) and a fourth lower edge plate (024); the first lower edge plate (021) and the second lower edge plate (022) are respectively provided with a bulge (07) which can be inserted into the limiting hole (06);
the bulges on the first lower edge plate (021) and the second lower edge plate (022) are respectively inserted into the two limiting holes (06), and the two limiting platforms (05) pass through the gaps formed by the first lower edge plate (021) and the second lower edge plate (022);
the third lower edge plate (023) and the fourth lower edge plate (024) are respectively covered on the outer surfaces of the first lower edge plate (021) and the second lower edge plate (022), and the third lower edge plate (023) and the fourth lower edge plate (024) are respectively fixed on the outer surfaces of the first lower edge plate (021) and the second lower edge plate (022) through riveting rivets (04); an engine connecting hole (03) is formed in the third lower edge plate (023);
the blade body (01), the lower edge plate assembly (02) and the riveting rivet (04) are all made of ceramic matrix composite materials;
integrally depositing a SiC ceramic matrix on the guide vane structure; depositing a SiC ceramic matrix at the reserved connection gap between the protrusion (07) and the limiting hole (06); and SiC ceramic matrixes are deposited at reserved connection gaps of the riveting rivet (04) and the third lower edge plate (023) and the fourth lower edge plate (024).
2. The aero engine ceramic matrix composite fixed guide vane structure of claim 1 wherein: set up breach (031) on third lower flange board (023), the inner edge of breach (031) and the outer fringe contact cooperation of two spacing platforms (05).
3. The aero engine ceramic matrix composite fixed guide vane structure of claim 1 or 2 wherein: the ceramic matrix composite is stacked by a plurality of layers of plain cloth; the blade body (01), the lower edge plate assembly (02) and the riveting rivet (04) are all formed by cutting along the stacking direction of the fiber cloth.
4. The forming of an aircraft engine ceramic matrix composite fixed guide vane structure as defined in claim 1, comprising the steps of:
preparing a blade body (01), a lower edge plate assembly (02) and riveting rivets (04) by using a ceramic matrix composite material;
secondly, the bulges (07) on the first lower edge plate (021) and the second lower edge plate (022) are respectively inserted into the two limiting holes (06), and the two limiting platforms (05) penetrate through the gaps formed by the first lower edge plate (021) and the second lower edge plate (022); positioning the first lower edge plate (021) and the second lower edge plate (022) is realized;
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 bulges (07);
step four, covering the surfaces of the first lower edge plate (021) and the second lower edge plate (022) with a third lower edge plate (023) and a fourth lower edge plate (024) respectively, and locking by riveting rivets (04); the inner edges of the notches (031) are in contact fit with the outer edges of the two limit platforms (05);
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 using SiC fibers, cutting the plain cloth into proper sizes according to the sizes of parts, 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 using 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 stacking direction of the fiber cloth to process the blade body (01), the lower edge plate assembly (02) and the riveting rivet (04);
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 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/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。
5. The forming of an aircraft engine ceramic matrix composite fixed guide vane structure as claimed in claim 4, 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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| CN112267917B (en) * | 2020-09-18 | 2022-09-23 | 中国航发四川燃气涡轮研究院 | Fiber preform and ceramic matrix composite turbine outer ring |
| CN112811921B (en) * | 2021-01-27 | 2022-10-25 | 巩义市泛锐熠辉复合材料有限公司 | Fiber-reinforced ceramic matrix composite heat-resistant plate and preparation method thereof |
| CN114427481B (en) * | 2022-01-06 | 2024-01-09 | 西北工业大学 | Ceramic matrix composite turbine rotor blade with internal cooling channels, mold and method of making |
| CN114109519A (en) * | 2022-01-28 | 2022-03-01 | 中国航发沈阳发动机研究所 | Guide blade based on metal base band ceramic mosaic block |
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