阅读说明：本技术 一种航空发动机用陶瓷基复合材料弹性密封件及制备方法 (Ceramic matrix composite elastic sealing element for aero-engine and preparation method thereof ) 是由 董宁 刘小冲 涂建勇 何江怡 刘持栋 孙肖坤 成来飞 于 2020-04-18 设计创作，主要内容包括：本发明涉及一种航空发动机用陶瓷基复合材料弹性密封件及制备方法,密封件为回转体或长条状,截面呈Ω形状,采用陶瓷基复合材料铆钉将本密封件铆接到航空发动机热端部件表面。Ω的上沿与另外一个密封面紧密接触。随着发动机环境温度升高和降低,以及发动机气流冲击、机械振动等环境条件变化,本密封件的开口尺寸L会被压缩或弹起,实现热端部件嵌套结构或对结面的有效密封。本发明制备的陶瓷基复合材料弹性密封件密度≥2.4g/cm<Sup>3</Sup>,材料孔隙率≤6％。1650K温度条件下强度保持率≥95％,密封材料基体开裂应力≥120MPa；承受应力≤100MPa时,保持线性回弹特性；1650K温度下使用寿命≥5000小时。(The invention relates to a ceramic matrix composite elastic sealing element for an aircraft engine and a preparation method thereof. The upper edge of Ω is in close contact with the other sealing surface. Along with the increase and decrease of the environmental temperature of the engine and the change of environmental conditions such as airflow impact, mechanical vibration and the like of the engine, the opening size L of the sealing element can be compressed or bounced, and the nesting structure of the hot end part or the effective sealing of the surface is realized. The inventionThe density of the prepared ceramic matrix composite elastic sealing element is more than or equal to 2.4g/cm 3 The porosity of the material is less than or equal to 6 percent. The strength retention rate under the temperature condition of 1650K is more than or equal to 95 percent, and the cracking stress of the sealing material matrix is more than or equal to 120 MPa; when the bearing stress is less than or equal to 100MPa, the linear rebound characteristic is kept; the service life is more than or equal to 5000 hours at 1650K.)

1. The utility model provides an aeroengine is with ceramic matrix composite elastic sealing member, is solid of revolution or rectangular form structure, its characterized in that: the section is in an omega shape, and the lower edge of the omega is fixed on the surface to be sealed by adopting the ceramic matrix composite rivet 2.

2. The ceramic matrix composite elastomeric seal for an aircraft engine of claim 1, wherein: the fibers include but are not limited to one or more of C fibers, SiC fibers, or boron nitride.

3. A method for preparing an elastic sealing element made of ceramic matrix composite material for an aeroengine according to claim 1 or 2, which is characterized by comprising the following steps:

step 1: weaving fibers into an elastomer preform with a 2D, 2.5D or 3D weaving structure, wherein the thickness of the preform is 4-6 mm;

step 2: fixing the prefabricated body between the omega-shaped graphite inner die and the outer die; the surface of the graphite mould is provided with a flow guide hole and an air guide hole in a ceramic process, and the hole diameter is 2-5 mm;

and step 3: placing a graphite mold with a fiber preform in CVD interface deposition equipment, wherein the deposition temperature of an interface layer is 500-1100 ℃, the vacuum of a deposition furnace is 3-50 kPa, 40-200L/min of Ar gas is used as protective gas, the flow rate of precursor gas of the interface layer is 100-500L/min, the deposition time is 20-50 h, and the thickness of the interface is controlled within the range of 500-700 mu m;

and 4, step 4: placing the preform after the deposition interface in CVI deposition equipment, and starting SiC matrix deposition: the deposition temperature of the SiC ceramic matrix is 1100-1400 ℃, and the deposition furnace is vacuumized to 20-50 kPa, 60-100L/min H₂The gas is used as carrier gas, the flow rate of the SiC ceramic matrix precursor gas is 100-500L/min, the single deposition time is 100-150 h, and the SiC matrix is densified and deposited for multiple times; when the density of the blank material is more than or equal to 1.3g/cm³Then, the graphite mold is removed, and the four steps are repeatedly circulated until the density is more than or equal to 2.0g/cm³Then finishing the preparation of a blank sealing piece;

and 5: processing a blank sealing element according to a target size, controlling the thickness of the sealing element to be 2-3 mm, and cleaning and drying after processing;

step 6: and (4) depositing a SiC anti-oxidation coating on the surface of the processed blank high-temperature elastic sealing element by adopting the CVI process of the step (4), and controlling the thickness to be 400-600 mu m to finish the preparation of the high-temperature elastic sealing element.

Technical Field

The invention belongs to the technical field of mechanical high-temperature sealing, and relates to a ceramic matrix composite elastic sealing element for an aircraft engine and a preparation method thereof.

Background

With the development of high-performance aircraft technology, the aerospace field has raised urgent demands for high-performance engines. Researches suggest that the improvement of the temperature of the turbine inlet gas becomes an effective way for improving the performance of the engine, and each country invests huge investment exploration and develops related technical examination and verification successively. The gas temperature before the turbine of the engine with the active push-to-weight ratio of 10 in the current developed country reaches 1850-1950K, the push-to-weight ratio of the fifth generation aircraft engine in the future is about 15-20, and the gas temperature before the turbine reaches 2200-2400K; the technical progress not only promotes the improvement of the temperature resistance of the materials for the engine, but also puts more severe requirements on the sealing performance among all hot-end components.

When the aircraft engine works, the gas leakage in the gas flow channel needs to be reduced or eliminated as much as possible so as to maximize the working efficiency of the gas. Gaps can be generated at the butt joint positions of the hot end parts of the engine due to the effects of temperature difference, air pressure difference, mechanical vibration and the like, or fixed gaps can be designed and reserved between the hot end parts to serve as air flow channels, and the gaps need to adopt high-reliability sealing to limit the direction of air flow or limit the area affected by the air flow, so that the normal operation of the aero-engine is guaranteed.

Chinese patent publication No. CN 103573416a of document 1 discloses a seal member. The seal disclosed in this patent is an annular or arcuate structure made from a high temperature alloy or a shape memory alloy. The patent does not provide the type of material of the seal ring and the temperature range of the seal ring.

The Chinese patent of 'patent publication No. CN 110318158A' in the document 2 discloses a high-temperature dynamic sealing material and a preparation method thereof. The patent relates to a preparation method of a temperature-resistant elastic sealing rope. The preparation method comprises the steps of elastic sealing rope forming, metal mesh pipe coating layer forming, fiber sheath coating layer forming, length cutting and end sealing treatment. The sealing rope disclosed by the patent has the characteristics of high temperature resistance, high compression performance and good resilience performance, has the advantage of adjustable compression force, and is suitable for high-temperature dynamic sealing and other parts. However, the temperature resistance limit of the elastic sealing rope disclosed by the patent is about 1100 ℃, and the diameter of the elastic sealing rope is small, so that the elastic sealing rope is inconvenient to fix and has poor adaptability to high-vibration environments.

In summary, the existing aircraft engine hot end seal has the following disadvantages:

(1) most of the existing high-temperature sealing elements are used for sealing conventional materials and conventional structures, and most of the sealing elements are of circular cross-section structures, so that the application range is limited greatly;

(2) the existing high-temperature sealing element usually adopts high-temperature alloy materials, inorganic fibers and the like. The material has high density and great weight, and the temperature resistance does not meet the development requirement of future high-performance engines.

Disclosure of Invention

Technical problem to be solved

In order to avoid the defects of the prior art, the invention provides the ceramic matrix composite elastic sealing element for the aero-engine and the preparation method thereof, which solve the problems of high density, poor temperature resistance, easy failure and the like of the conventional high-temperature elastic sealing element and meet the development requirement of high-temperature sealing of hot end components of the high-performance aero-engine in the future.

Technical scheme

The utility model provides an aeroengine is with ceramic matrix composite elastic sealing member, is solid of revolution or rectangular form structure, its characterized in that: the section is in an omega shape, and the lower edge of the omega is fixed on the surface to be sealed by adopting the ceramic matrix composite rivet 2.

The fibers include but are not limited to one or more of C fibers, SiC fibers, or boron nitride.

The preparation method of the ceramic matrix composite elastic sealing element for the aero-engine is characterized by comprising the following steps:

step 1: weaving fibers into an elastomer preform with a 2D, 2.5D or 3D weaving structure, wherein the thickness of the preform is 4-6 mm;

step 2: fixing the prefabricated body between the omega-shaped graphite inner die and the outer die; the surface of the graphite mould is provided with a flow guide hole and an air guide hole in a ceramic process, and the hole diameter is 2-5 mm;

and step 3: placing a graphite mold with a fiber preform in CVD interface deposition equipment, wherein the deposition temperature of an interface layer is 500-1100 ℃, the vacuum of a deposition furnace is 3-50 kPa, 40-200L/min of Ar gas is used as protective gas, the flow rate of precursor gas of the interface layer is 100-500L/min, the deposition time is 20-50 h, and the thickness of the interface is controlled within the range of 500-700 mu m;

and 4, step 4: placing the preform after the deposition interface in CVI deposition equipment, and starting SiC matrix deposition: the deposition temperature of the SiC ceramic matrix is 1100-1400 ℃, and the deposition furnace is vacuumized to 20-50 kPa, 60-100L/min H₂The gas is used as carrier gas, the flow rate of the SiC ceramic matrix precursor gas is 100-500L/min, the single deposition time is 100-150 h, and the SiC matrix is densified and deposited for multiple times; when the density of the blank material is more than or equal to 1.3g/cm³Then, the graphite mold is removed, and the four steps are repeatedly circulated until the density is more than or equal to 2.0g/cm³Then finishing the preparation of a blank sealing piece;

and 5: processing a blank sealing element according to a target size, controlling the thickness of the sealing element to be 2-3 mm, and cleaning and drying after processing;

step 6: depositing a SiC anti-oxidation coating on the surface of the processed blank high-temperature elastic sealing element by adopting the CVI process of the step 4, controlling the thickness to be 400-600 mu m, and finishing the preparation of the high-temperature elastic sealing element

Advantageous effects

The cross section of the elastic sealing element is shown in figure 1, the sealing element 1 is a rotary body or a strip shape, the cross section is in an omega shape, and the sealing element 1 is riveted to the surface of a hot end component 3 of the aero-engine by adopting a rivet 2 made of a ceramic matrix composite material. The upper edge of Ω is in close contact with the other sealing surface. Along with the increase and decrease of the environmental temperature of the engine and the change of the environmental conditions such as the air flow impact, the mechanical vibration and the like of the engine, the opening size L of the sealing element 1 can be compressed or bounced, and the nesting structure of the hot end part or the effective sealing of the surface is realized.

The invention can cover the mechanical seal between the ceramic matrix composite hot end parts under the environment of RT-1700 ℃, and meets the requirements of dynamic, long-service life and high-reliability sealing connection of precise mechanical parts in the aerospace field. The density of the elastic sealing element made of the ceramic matrix composite material prepared by the invention is more than or equal to 2.4g/cm³The porosity of the material is less than or equal to 6 percent. The strength retention rate under the temperature condition of 1650K is more than or equal to 95 percent, and the cracking stress of the sealing material matrix is more than or equal to 120 MPa; when the bearing stress is less than or equal to 100MPa, the linear rebound characteristic is kept; the service life is more than or equal to 5000 hours at 1650K.

Drawings

FIG. 1 is a schematic view of a high temperature elastomeric seal interface and seal arrangement;

the reference numbers in the figures are: the method comprises the following steps of 1-a schematic cross-sectional view of the ceramic matrix composite elastic sealing element, 2-a schematic cross-sectional view of a high-temperature-resistant riveting structure and 3-a schematic cross-sectional view of movable parts in an engine.

Detailed Description

The invention will now be further described with reference to the following examples and drawings:

examples 1

Step 1, preparing an elastomer preform by using the C fiber, wherein the preform structure is a 2D structure, and the thickness of the preform is 4 mm.

And 2, placing the prefabricated body between the graphite inner die and the graphite outer die, and compressing. The surface of the graphite mould is punched, and the diameter of the hole is 2 mm.

Step 3, placing the fiber preform with the mold in CVD interface deposition equipment, wherein the deposition temperature of an interface layer is 500-1100 ℃, the vacuum of a deposition furnace is 3-50 kPa, 90L/min of Ar gas is used as protective gas, boron trichloride is used as precursor gas of the interface layer, the flow rate of the boron trichloride gas is 140L/min, the deposition time is 30h, and the thickness of the interface is controlled at 600 mu m;

and 4, placing the preform with the deposited interface in CVI deposition equipment, and starting SiC matrix deposition. The deposition temperature of the SiC ceramic matrix is 1100-1400 ℃, and the deposition furnace is vacuumized to 20-50 kPa, 80L/min H₂Gas is used as carrier gas, trichloromethylsilane is used as a SiC ceramic matrix precursor, the gas flow is 200L/min, the single deposition time is 100h, and the SiC matrix is deposited for multiple times; when the density of the blank material is more than or equal to 1.3g/cm³Then, removing the graphite mold; repeating the step 4 until the density is more than or equal to 2.0g/cm³And then finishing the preparation of the blank sealing member.

And 5, machining the blank sealing element according to the target size by adopting a diamond cutter, controlling the thickness of the sealing element to be 2mm, cleaning and drying after machining, and transferring to the next procedure.

And 6, depositing a SiC anti-oxidation coating on the surface of the processed blank high-temperature elastic sealing element by adopting the CVI process in the step 4, and controlling the thickness to be 400 mu m to finish the preparation of the high-temperature elastic sealing element.

And 7, riveting the prepared high-temperature elastic sealing element at the part of the hot end component of the engine, which needs to be sealed, by adopting the ceramic matrix composite rivet, and finishing the installation of the high-temperature elastic sealing element.

The density of the elastic seal obtained in this example was 2.0g/cm³The porosity of the material is less than 6%. The strength retention rate under 1650K condition is more than or equal to 96 percent, and the cracking stress of the matrix is 160 MPa; when the bearing stress is less than or equal to 100MPa, the linear rebound is kept, and the high-temperature service life of 1650K is more than or equal to 6000 hours.

EXAMPLES example 2

Step 1, preparing an elastomer preform by using SiC fibers, wherein the preform structure is a 2.5D structure, and the thickness of the preform is 5 mm.

And 2, placing the prefabricated body between the graphite inner die and the graphite outer die, and compressing. The surface of the graphite mould is punched, and the diameter of the hole is 4 mm.

Step 3, placing the fiber preform with the mold in CVD interface deposition equipment, wherein the deposition temperature of an interface layer is 500-1100 ℃, the vacuum of a deposition furnace is 3-50 kPa, 180L/min of Ar gas is used as protective gas, boron trichloride is used as precursor gas of the interface layer, the flow rate of the boron trichloride gas is 300L/min, the deposition time is 40h, and the thickness of the interface is controlled at 650 mu m;

and 4, placing the preform with the deposited interface in CVI deposition equipment, and starting SiC matrix deposition. The deposition temperature of the SiC ceramic matrix is 1100-1400 ℃, and the deposition furnace is vacuumized to 20-50 kPa, 90L/min H₂Gas is used as carrier gas, trichloromethylsilane is used as a precursor of the SiC ceramic matrix, the gas flow is 400L/min, the single deposition time is 120h, and the SiC matrix is deposited for multiple times; when the density of the blank material is more than or equal to 1.3g/cm³Then, removing the graphite mold; repeating the steps until the density is more than or equal to 2.0g/cm³And then finishing the preparation of the blank sealing member.

And 5, machining the blank sealing element according to the target size by adopting a diamond cutter, controlling the thickness of the sealing element to be 2.5mm, cleaning and drying after machining, and transferring to the next procedure.

And 6, depositing a SiC anti-oxidation coating on the surface of the processed blank high-temperature elastic sealing element by adopting the CVI process in the step 1, and controlling the thickness to be 500 mu m to finish the preparation of the high-temperature elastic sealing element.

And 7, riveting the prepared high-temperature elastic sealing element at the part of the hot end component of the engine, which needs to be sealed, by adopting the ceramic matrix composite rivet, and finishing the installation of the high-temperature elastic sealing element.

The density of the elastic seal obtained in this example was 2.5g/cm³The porosity of the material is less than 5%. The strength retention rate under 1650K condition is more than or equal to 98 percent, and the cracking stress of the matrix is 170 MPa; when the bearing stress is less than or equal to 100MPa, the linear rebound can be kept, and the high-temperature service life of 1650K is more than or equal to 10000 hours.

EXAMPLE 3

Step 1, preparing an elastomer preform by adopting fibers such as boron nitride and the like, wherein the preform is of a 3D structure, and the thickness of the preform is 6 mm.

And 2, placing the prefabricated body between the graphite inner die and the graphite outer die, and compressing. The surface of the graphite mould is punched, and the diameter of the hole is 5 mm.

Step 3, placing the fiber preform with the mold in CVD interface deposition equipment, wherein the deposition temperature of an interface layer is 500-1100 ℃, the vacuum of a deposition furnace is 3-50 kPa, 200L/min of Ar gas is used as protective gas, boron trichloride is used as precursor gas of the interface layer, the flow rate of the boron trichloride gas is 450L/min, the deposition time is 50h, and the thickness of the interface is controlled at 700 mu m;

and 4, placing the preform with the deposited interface in CVI deposition equipment, and starting SiC matrix deposition. The deposition temperature of the SiC ceramic matrix is 1100-1400 ℃, and the deposition furnace is vacuumized to 20-50 kPa, 100L/min H₂Gas is used as carrier gas, trichloromethylsilane is used as a SiC ceramic matrix precursor, the gas flow is 500L/min, the single deposition time is 150h, and SiC matrix is deposited for multiple times; when the density of the blank material is more than or equal to 1.3g/cm³Then, removing the graphite mold; repeating the steps until the density is more than or equal to 2.0g/cm³And then finishing the preparation of the blank sealing member.

And 5, machining the blank sealing element according to the target size by adopting a diamond cutter, controlling the thickness of the sealing element to be 3mm, cleaning and drying after machining, and transferring to the next procedure.

And 6, depositing a SiC anti-oxidation coating on the surface of the processed blank high-temperature elastic sealing element by adopting the CVI process in the step 4, and controlling the thickness to be 600 mu m to finish the preparation of the high-temperature elastic sealing element.

And 7, riveting the prepared high-temperature elastic sealing element at the part of the hot end component of the engine, which needs to be sealed, by adopting the ceramic matrix composite rivet, and finishing the installation of the high-temperature elastic sealing element.

The density of the elastic seal obtained in this example was 2.5g/cm³The porosity of the material is less than 5%. The strength retention rate under 1650K condition is more than or equal to 93 percent, and the cracking stress of the matrix is 150 MPa; when the bearing stress is less than or equal to 100MPa, the linear rebound is kept, and the high-temperature service life of 1650K is more than or equal to 15000 hours.