阅读说明：本技术 一种陶瓷基复合材料弹性密封件及其成型工艺 (Ceramic matrix composite elastic sealing element and forming process thereof ) 是由 涂建勇 王佳民 刘梦珠 王文红 成来飞 于 2020-03-31 设计创作，主要内容包括：本发明属于航空发动机热端结构件密封技术领域,具体涉及一种陶瓷基复合材料弹性密封件及其成型工艺,解决现有高温弹性密封件普遍存在耐温性能不佳、自重较高等缺陷,包括长条状或环形的弹性密封件本体,弹性密封件本体的材质为陶瓷基复合材料；陶瓷基复合材料为C/SiC陶瓷基复合材料或SiC/SiC陶瓷基复合材料；弹性密封件本体的横截面形状为N形,相邻两个壁面通过弧面圆滑过渡连接；在相对且平行的两个壁面的其中一个壁面上开设有连接孔。可以实现1200K～1700K高温环境下陶瓷基复合材料热端部件之间的机械密封,满足航空航天领域精密机械的长寿命、高可靠密封连接。(The invention belongs to the technical field of sealing of hot end structural members of aero-engines, and particularly relates to an elastic sealing element made of ceramic matrix composite and a forming process thereof, which solve the defects of poor temperature resistance, high self weight and the like of the conventional high-temperature elastic sealing element, wherein the conventional high-temperature elastic sealing element comprises a strip-shaped or annular elastic sealing element body, and the elastic sealing element body is made of the ceramic matrix composite; the ceramic matrix composite material is a C/SiC ceramic matrix composite material or a SiC/SiC ceramic matrix composite material; the cross section of the elastic sealing element body is N-shaped, and two adjacent wall surfaces are in smooth transition connection through an arc surface; one of the two opposite and parallel wall surfaces is provided with a connecting hole. The mechanical seal between the ceramic matrix composite hot end parts under the high-temperature environment of 1200-1700K can be realized, and the requirements of long service life and high reliability sealing connection of precision machinery in the aerospace field are met.)

1. An elastomeric seal of ceramic matrix composite, comprising: an elastomeric seal body comprising an elongated or annular shape; the elastic sealing element body is made of a ceramic matrix composite material; the ceramic matrix composite material is a C/SiC ceramic matrix composite material or a SiC/SiC ceramic matrix composite material;

the cross section of the elastic sealing element body is N-shaped, and two adjacent wall surfaces are in smooth transition connection through an arc surface; one of the two opposite and parallel wall surfaces is provided with a connecting hole.

2. A seal structure of an engine component, characterized in that: comprising a connection member (2) and the ceramic matrix composite elastomeric seal of claim 1;

the elastic sealing element of the ceramic matrix composite material is positioned between two engine components to be sealed, a first wall surface (21) of two opposite parallel wall surfaces is abutted against the sealing surface of one engine component to be sealed, and the outer wall of a second wall surface (22) is abutted against the sealing surface of the other engine component to be sealed;

the connecting piece (2) passes through the connecting hole of the first wall surface (21) to fix the elastic sealing piece of the ceramic matrix composite material on the sealing surface of one engine component to be sealed.

3. The seal structure of an engine component according to claim 2, characterized in that: the connecting piece (2) is made of a ceramic matrix composite material.

4. A forming process of an elastic sealing element made of ceramic matrix composite materials is characterized in that: the method comprises the following steps:

step 1, preparing a blank ceramic matrix composite;

step 1.1, preparing a fiber preform;

weaving C fibers or SiC fibers into a 2D or 2.5D fiber preform;

step 1.2, preparing an interface layer;

placing the fiber preform fixedly formed in the step 1.1 into a deposition furnace, and preparing an interface layer;

step 1.3, depositing a SiC ceramic matrix;

placing the product prepared in the step 1.2 in CVI deposition equipment, and depositing a SiC ceramic matrix on the product by adopting a CVI process to obtain a blank ceramic matrix composite;

step 2, processing the blank ceramic matrix composite;

placing the blank ceramic matrix composite material prepared in the step 1 on processing equipment, cutting the blank ceramic matrix composite material into an N shape along the thickness direction by using a processing cutter, and forming a connecting hole on one of two opposite and parallel wall surfaces;

step 3, obtaining the ceramic matrix composite elastic sealing element;

and (3) placing the product processed in the step (2) in a chemical vapor deposition furnace (CVI), and depositing a SiC ceramic matrix coating on the product by adopting a CVI process, wherein the thickness of the coating is controlled to be 20-200 mu m, so as to obtain the ceramic matrix composite elastic sealing element.

5. The process of forming a ceramic matrix composite elastomeric seal according to claim 4, wherein the process conditions for preparing the interface layer in step 1.2 are:

the deposition temperature of the interface layer is 500-1000 ℃, the deposition furnace is vacuumized to 3-50 kPa, 40-200L/min 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 layer is 500-700 mu m.

6. The forming process of the ceramic matrix composite elastic sealing element according to claim 5, wherein the process conditions for preparing the interface layer in the step 1.2 are that the deposition temperature of the interface layer is 800 ℃, the deposition furnace is vacuumized to 40kPa, 200L/min Ar gas is used as protective gas, the flow rate of the precursor gas of the boron trichloride interface layer is 500L/min, the deposition time is 50h, and the interface thickness is controlled at 700 μm.

7. The process for forming an elastomeric seal of ceramic matrix composite according to claim 4, wherein the process conditions for depositing SiC ceramic matrix on the product by CVI process in step 1.3 and step 3 are:

the temperature is 1100-1400 ℃, the chemical vapor deposition furnace is vacuumized to 20-50 kPa, H of 60-100L/min₂The gas is used as carrier gas, the SiC ceramic precursor is trichloromethylsilane, the Ar gas is used as diluent gas, the flow rate of the SiC ceramic precursor is 100-500L/min, and the single deposition time is 100-150 h.

8. The process for forming an elastomeric seal of ceramic matrix composite according to claim 7, wherein the process conditions for depositing SiC ceramic matrix on the product by CVI process in step 1.3 and step 3 are:

the temperature is 1300 ℃, the deposition furnace is vacuumized to 50kPa, H of 100L/min₂The gas is used as carrier gas, the flow rate of the MTS gas of the SiC ceramic matrix precursor is 500L/min, and the single deposition time is 150 h.

9. The process of forming a ceramic matrix composite elastomeric seal according to claim 7, wherein: step 1.3, the CVI process is repeated for multiple times until the blank pottery is obtainedThe density of the ceramic-based composite material is more than or equal to 2.0g/cm³。

10. The process of forming a ceramic matrix composite elastomeric seal according to claim 7, wherein: and 2, adopting a common multi-axis numerical control machine tool as processing equipment, and adopting cubic boron nitride or diamond as a processing cutter.

Technical Field

The invention belongs to the technical field of sealing of hot end structural members of aero-engines, and particularly relates to a radial self-tightening type elastic sealing member made of a ceramic matrix composite material and a forming process of the sealing member.

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 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 aircraft engine in the future can reach about 15-20, the gas temperature before the turbine can reach 2200-2400K, and the temperature is far beyond the heat-resisting limit of the current turbine and turbine front-end component material.

Meanwhile, with the continuous improvement of the performance requirements of modern aircraft engines, the problem of sealing between engine components has become a research hotspot in the field of engines. When the aircraft engine works, the air pressure leakage of the air flow channel needs to be reduced or eliminated as much as possible so as to realize the maximization of the work done by the air. At the butt joint part of the hot end component, the gaps at low temperature and high temperature are different, and in order to ensure the sealing action under the working conditions of low temperature and high temperature, an elastic sealing element is usually selected to bear the impact of vibration, high temperature difference and high pressure difference and severe deformation caused by the environment alternation.

Chinese patent publication No. CN102537350A of document 1 discloses a seal ring and an aircraft engine having the same. The sealing ring disclosed in this patent is annular and has a "W" shaped cross-section with two outer walls on the outside and two inner walls on the inside. The patent does not provide the type of material of the seal ring and the temperature range of the seal ring.

The chinese invention patent of document 2 "patent publication No. CN 203560450U" discloses a metal seal ring for an aircraft engine. The sealing ring adopts a high-hardness nickel-based alloy material as a ring body substrate, the section of the ring body structure is W-shaped, and the sealing ring can fully ensure the reliable work of the system work by virtue of the self-sealing effect of the environmental pressure. However, the sealing ring is made of traditional high-temperature nickel-based alloy, the temperature resistance of the sealing ring does not exceed 1150 ℃, and the performance requirement of a high-performance aeroengine can still not be met obviously.

In summary, the existing sealing technology for the hot end part of the aircraft engine has the following defects:

(1) the design idea of the existing high-temperature sealing ring still aims at the conventional materials, does not consider the problems of other new processes, selection and forming of new materials and the like, and has the problem of poor temperature resistance;

(2) most of the existing high-temperature sealing rings are made of high-temperature alloy materials, the material density is high, the self weight of the sealing structure is high, and meanwhile, the temperature resistance performance of the sealing structure cannot meet the requirement of a future high-performance engine.

(3) And the problems of a high-temperature sealing ring connecting mode, a fixing mode and the like are not considered in the conventional patent.

Disclosure of Invention

The invention discloses a ceramic matrix composite elastic sealing element and a forming process thereof, aiming at the defects of poor temperature resistance, high self weight and the like of the existing high-temperature elastic sealing element generally so as to meet the performance development requirement of a future high-performance aircraft engine.

The technical scheme of the invention is to provide an elastic sealing element made of ceramic matrix composite, which is characterized in that: the elastic sealing element comprises a strip-shaped or annular elastic sealing element body, wherein the elastic sealing element body is made of a ceramic matrix composite material; the ceramic matrix composite material is a C/SiC ceramic matrix composite material or a SiC/SiC ceramic matrix composite material;

the cross section of the elastic sealing element body is N-shaped, and two adjacent wall surfaces are in smooth transition connection through an arc surface; one of the two opposite and parallel wall surfaces is provided with a connecting hole.

The invention also provides a sealing structure of the engine component, which is characterized in that: comprises a connecting piece and the ceramic matrix composite elastic sealing element;

the ceramic matrix composite elastic sealing element is positioned between two engine components to be sealed, a first wall surface of two relatively parallel wall surfaces is abutted against a sealing surface of one engine component to be sealed, and the outer wall of a second wall surface is abutted against a sealing surface of the other engine component to be sealed;

the connecting piece penetrates through the connecting hole on the first wall surface to fix the elastic sealing piece made of the ceramic matrix composite material on the sealing surface of one engine component to be sealed.

In order to further improve the temperature resistance, the connecting piece is made of a ceramic matrix composite material.

The invention also provides a forming process of the ceramic matrix composite elastic sealing element, which comprises the following steps:

step 1, preparing a blank ceramic matrix composite;

step 1.1, preparing a fiber preform;

weaving C fibers or SiC fibers into a 2D or 2.5D fiber preform;

step 1.2, preparing an interface layer;

placing the fiber preform fixedly formed in the step 1.1 into a deposition furnace, and preparing an interface layer;

step 1.3, depositing a SiC ceramic matrix;

placing the product prepared in the step 1.2 in CVI deposition equipment, and depositing a SiC ceramic matrix on the product by adopting a CVI process to obtain a blank ceramic matrix composite;

step 2, processing the blank ceramic matrix composite;

placing the blank ceramic matrix composite prepared in the step 1 on processing equipment, cutting the blank ceramic matrix composite along the thickness direction by using a processing cutter, processing the appearance structure of the ceramic matrix composite elastic sealing element, the connecting hole and the connecting piece into an N shape, and forming the connecting hole on one of two opposite and parallel wall surfaces;

step 3, obtaining the ceramic matrix composite elastic sealing element;

and (3) placing the product processed in the step (2) in a chemical vapor deposition furnace (CVI), and depositing a SiC ceramic matrix coating on the product by adopting a CVI process, wherein the thickness of the coating is controlled to be 20-200 mu m, so as to obtain the ceramic matrix composite elastic sealing element.

Further, the process conditions for preparing the interface layer in the step 1.2 are as follows:

the deposition temperature of the interface layer is 500-1000 ℃, the deposition furnace is vacuumized to 3-50 kPa, 40-200L/min 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 layer is 500-700 mu m.

Further, the preferred process conditions for preparing the interface layer in step 1.2 are as follows:

the deposition temperature of the interface layer is 800 ℃, the deposition furnace is vacuumized to 40kPa, 200L/min Ar gas is used as protective gas, the flow rate of the precursor gas of the boron trichloride interface layer is 500L/min, the deposition time is 50h, and the interface thickness is controlled at 700 mu m.

Further, the process conditions for depositing the SiC ceramic matrix on the product by adopting the CVI process in step 1.3 and step 3 are as follows:

the temperature is 1100-1400 ℃, the chemical vapor deposition furnace is vacuumized to 20-50 kPa, H of 60-100L/min₂The gas is used as carrier gas, the SiC ceramic precursor is trichloromethylsilane, the Ar gas is used as diluent gas, the flow rate of the SiC ceramic precursor is 100-500L/min, and the single deposition time is 100-150 h.

Further, the preferred process conditions for depositing the SiC ceramic matrix on the product by the CVI process in step 1.3 and step 3 are as follows:

the temperature is 1300 ℃, the deposition furnace is vacuumized to 50kPa, H of 100L/min₂The gas is used as carrier gas, the flow rate of the MTS gas of the SiC ceramic matrix precursor is 500L/min, and the single deposition time is 150 h.

Further, repeating the CVI process for a plurality of cycles in step 1.3 until the density of the green ceramic matrix composite material is greater than or equal to 2.0g/cm³。

Further, in the step 2, 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. the invention can realize the mechanical sealing between the ceramic matrix composite hot end components in the high temperature environment of 1200K-1700K, and can meet the requirements of long service life and high reliability sealing connection of precision machinery in the aerospace field.

2. The density of the elastic sealing element made of the ceramic matrix composite material is 2.0-2.5 g/cm³. The strength retention rate is more than or equal to 95% under the condition of 1600K, the porosity of the material is less than 5%, and the good resilience performance can be kept at high temperature.

3. The elastic sealing element with the N-shaped section has the advantages of simple preparation process, large elastic compressible capacity and high installation reliability, and the application environment mainly comprises the following steps:

a. mechanical seals between large high temperature components; b. the two high-temperature components to be sealed are mechanical seals with different materials and large difference of expansion coefficients (the environment is the same, and the materials are different); c. two high-temperature components needing to be sealed are mechanical seals (same materials and different environments) with large temperature difference.

Drawings

FIG. 1 is a schematic view of the sealing principle and its cross-sectional configuration of a ceramic matrix composite elastomeric seal;

the reference numbers in the figures are: 1-ceramic matrix composite elastic sealing element, 2-connecting element, 21-first wall surface, 22-second wall surface, and 3-sealing surface of engine component to be sealed.

FIG. 2 is a schematic view of the sealing of the ceramic matrix composite elastomeric seal as the size of the opening is reduced;

Detailed Description

The section of the elastic sealing element made of the ceramic matrix composite material is shown in figure 1, the elastic sealing element 1 made of the ceramic matrix composite material can be an annular or strip-shaped part with an N-shaped section, two adjacent wall surfaces are in smooth transition connection through an arc surface, the two opposite and parallel wall surfaces are defined as a first wall surface 21 and a second wall surface 22, a connecting hole is formed in the first wall surface 21, the elastic sealing element 1 made of the ceramic matrix composite material is riveted on a sealing surface of one engine component to be sealed by adopting a connecting piece 2, the connecting piece 2 can be a SiC/SiC riveting piece, and by combining the figure 1 and the figure 2, along with temperature rise and temperature fall and mechanical vibration inside an engine, the opening size L of the elastic sealing element made of the ceramic matrix composite material can be reduced or enlarged, and the second wall surface 22 is always kept in close contact with the sealing surface 3 of the engine component to be sealed, so that effective sealing is realized.

Preparing a ceramic matrix composite elastomeric seal by:

step one, preparing a fiber preform by using C fibers or SiC fibers.

And secondly, placing the fiber preform with the mold in a deposition furnace, wherein the deposition temperature of an interface layer is 500-1000 ℃, the deposition furnace is vacuumized to 3-50 kPa, 40-200L/min Ar gas is used as a protective gas, the flow rate of a 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.

Step three, placing the preform with the deposited interface in CVI deposition equipment, and starting SiC matrix deposition, wherein the deposition temperature of the SiC ceramic matrix is 1100-1400 ℃, and the deposition furnace is vacuumized to 20-50 kPa and H of 60-100L/min₂The 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, the SiC matrix is densified and deposited for many times, and when the density of the blank material is more than or equal to 2.0g/cm³And then transferring the prepared blank ceramic matrix composite material to the next working procedure.

And step four, machining the blank ceramic matrix composite material according to the target size by adopting a cubic boron nitride or diamond cutter, and finishing the preparation of the blank high-temperature elastic sealing element.

And step five, depositing an anti-oxidation SiC coating on the surface of the blank high-temperature elastic sealing element by adopting a process of step three, and controlling the thickness of the coating to be 20-200 mu m to finish the preparation of the high-temperature elastic sealing element.

And step six, riveting the prepared high-temperature elastic sealing element on a part needing to be sealed of the hot end component of the engine by adopting the ceramic matrix composite rivet to finish the installation of the high-temperature elastic sealing element.

The present invention is further described with reference to the following examples.