Ceramic-based turbine rotor of aero-engine and preparation process of ceramic-based turbine rotor

Abstract

The invention discloses a ceramic-based turbine rotor of an aero-engine and a preparation process of the ceramic-based turbine rotor, and relates to the technical fields of engine structure design, material application, forming processes and the like. The fiber prefabricated body is formed by stacking common 2D plain fiber cloth from bottom to top, the included angle between fibers of an upper laying layer and fibers of a lower laying layer is fixed to be 30-45 degrees, after the laying layers reach the thickness, the fibers of all the layers are sewn and connected through a puncture method, forming of the turbine rotor prefabricated body is completed, and the fiber prefabricated body is of an integral leaf disc structure; compared with the mode that the independent ceramic-based turbine blades and the turbine disc are connected through tenon grooves, the number of parts is reduced, the part structure is simplified, and the reliability of an engine is improved.

Description

Technical Field

[0001] The present invention relates to the technical fields of engine structure design, material application, molding technology, etc., and specifically to a ceramic-based turbine rotor for an aero-engine and a preparation process thereof. Background Art

[0002] An aircraft engine is a highly complex and sophisticated thermal machine. As the core hot end component of an aircraft engine, the turbine rotor converts the thermal energy of high-temperature, high-pressure combustion gas into mechanical energy to achieve work, and has particularly stringent requirements on temperature resistance. During operation, the turbine rotor needs to withstand the combined effects of high temperature and high pressure, temperature and centrifugal alternating loads, vibration loads, thermal stress, combustion gas corrosion and high-temperature oxidation.

[0003] At present, the alloy material microstructure is optimized and the comprehensive performance of the alloy is improved by applying alloying theory, heat treatment, smelting and casting technology. Improvement measures are mainly taken in the structural design of turbine blades and the surface protection of hot end components. For example, the turbine cooling technology is improved by adopting a double-wall structure for turbine blades, optimizing and improving the structural design of the air inlet passage, and improving manufacturing. A thermal barrier coating with low thermal conductivity, high melting point, high temperature phase structure and stable chemistry is applied to the surface of the alloy substrate to improve its high-temperature oxidation and heat corrosion resistance, so as to reduce the temperature of the base material, improve the temperature bearing capacity of the turbine blades, increase the turbine inlet temperature, and improve the engine performance.

[0004] Currently, the demand for increasing the temperature of the air inlet to the turbine is greater than the temperature increase space that the turbine blades can withstand. The turbine inlet temperature has gradually approached the melting point of high-temperature alloys. Adding refractory metal elements to high-temperature alloys faces many bottleneck problems that need to be tackled. Summary of the invention

[0005] Therefore, in order to solve the above-mentioned deficiencies, the present invention provides an aero-engine ceramic-based turbine rotor and a preparation process thereof.

[0006] The present invention is implemented by constructing an aircraft engine ceramic-based turbine rotor and a preparation process thereof, comprising the following steps;

[0007] S1: Fiber preform: Ordinary 2D plain fiber cloth is stacked from bottom to top, and the angle between the upper and lower layers of fiber is fixed at 30° to 45°. After the layers reach the required thickness, the fibers of each layer are sutured and connected by puncture to complete the formation of the turbine rotor preform;

[0008] S2: The turbine rotor uses a SiC substrate, and the turbine rotor preform is placed in a reaction chamber at 900-1100°C and 4-10KPa;

[0009] S3: mixing reaction gas and carrier gas into the reaction chamber;

[0010] S4: The pressure above the preform in the reaction chamber is atmospheric pressure, and the precursor pressure is 3.5 MPa higher than atmospheric pressure.

[0011] In a feasible implementation manner, the fiber preform is an integral turbine blade disk structure.

[0012] In a feasible implementation, the reaction gas uses a gaseous precursor with a Si:C chemical ratio of 1:1 as the deposited SiC, the gaseous precursor is composed of trichloromethylsilane, H2 is used as a carrier gas, and is mixed into the mixture according to H2 / trichloromethylsilane 4-8.

[0013] The present invention has the following advantages: The present invention provides an aircraft engine ceramic-based turbine rotor and a preparation process thereof through improvement, which has the following improvements compared with similar equipment:

[0014] The invention discloses an aero-engine ceramic-based turbine rotor and a preparation process thereof, wherein the fiber preform adopts an integral blade disk structure, which reduces the number of parts, simplifies the part structure, and improves the engine reliability compared to a single ceramic-based turbine blade connected to a turbine disk through a falcon groove.

[0015] The invention discloses a ceramic-based turbine rotor for an aero-engine and a preparation process thereof. The ceramic-based composite material has a low density. Compared with turbines made of traditional high-temperature alloy materials, the turbine made from the rotor achieves lightweight components and greatly reduces the weight of the engine, thereby greatly improving the thrust-to-weight ratio.

[0016] The invention discloses an aero-engine ceramic-based turbine rotor and a preparation process thereof. The ceramic-based composite material has strong stability and corrosion resistance, reduces the necessary measures for conventional coating treatment to improve the corrosion resistance of blades, and reduces the process difficulty of coating application. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 It is a schematic structural diagram of a first mode of lamination of a turbine rotor preform of the present invention;

[0018] Figure 2 It is a schematic structural diagram of a second method of laminating a turbine rotor preform according to the present invention;

[0019] Figure 3 It is a schematic structural diagram of the third method of lamination of turbine rotor preforms of the present invention. DETAILED DESCRIPTION

[0020] The following will be combined with the attached Figure 1-Figure 3The present invention is described in detail, and the technical solutions in the embodiments of the present invention are made clear. The complete description is obvious. The described embodiments are only part of the embodiments of the present invention, not all of them. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0021] It should be noted that when a component is referred to as being "fixed to" another component, it may be directly on the other component or there may also be a component centered. When a component is considered to be "connected to" another component, it may be directly connected to the other component or there may also be a component centered. When a component is considered to be "set on" another component, it may be directly set on the other component or there may also be a component centered. The terms "vertical", "horizontal", "left", "right" and similar expressions used herein are for illustrative purposes only.

[0022] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as those commonly understood by those skilled in the art of the present invention. The terms used herein in the specification of the present invention are only for the purpose of describing specific embodiments and are not intended to limit the present invention. The term "and / or" used herein includes any and all combinations of one or more related listed items.

[0023] Embodiment 1;

[0024] See also Figures 1 to 3 , an aero-engine ceramic-based turbine rotor and a preparation process thereof of the present invention comprises the following steps;

[0025] S1: Fiber preform: Ordinary 2D plain fiber cloth is stacked from bottom to top, and the angle between the upper and lower layers of fiber is fixed at 30°. After the layers reach the required thickness, the fiber layers are sutured and connected by puncture to complete the formation of the turbine rotor preform;

[0026] S2: Ceramic preparation: The turbine rotor uses SiC substrate, and the turbine rotor preform is placed in a reaction chamber at 900℃ and 4KPa;

[0027] S3: mixing reaction gas and carrier gas into the reaction chamber;

[0028] S4: The pressure above the preform in the reaction chamber is atmospheric pressure, and the precursor pressure is 3.5 MPa higher than atmospheric pressure. The pressure difference causes the gaseous precursor to flow and diffuse around and inside the turbine rotor fiber preform, prompting the precursor gas to undergo a chemical reaction in the pores, depositing to form a ceramic matrix, filling the pores and forming a dense composite material.

[0029] Specifically, the fiber preform is an integral turbine blade disk structure, that is, 33 turbine working blades and a turbine disk are an integral structure, and the blades and other complex surfaces are three-dimensionally woven.

[0030] Specifically, the reaction gas uses a gaseous precursor with a chemical ratio of Si:C of 1:1 as the deposited SiC, the gaseous precursor is composed of trichloromethylsilane, H2 is used as a carrier gas, and is mixed into the reaction gas at a ratio of H2 / trichloromethylsilane of 4 to 8.

[0031] Embodiment 2:

[0032] See also Figures 1 to 3 , an aero-engine ceramic-based turbine rotor and a preparation process thereof of the present invention, this embodiment also includes;

[0033] The steps include:

[0034] S1: Fiber preform: Ordinary 2D plain fiber cloth is stacked from bottom to top, and the angle between the upper and lower layers of fiber is fixed at 45°. After the layers reach the required thickness, the fiber layers are sutured and connected by puncture to complete the formation of the turbine rotor preform;

[0035] S2: Ceramic preparation: The turbine rotor uses SiC substrate, and the turbine rotor preform is placed in a reaction chamber at 1000°C and 10KPa;

[0036] S3: mixing reaction gas and carrier gas into the reaction chamber;

[0037] S4: The pressure above the preform in the reaction chamber is atmospheric pressure, and the precursor pressure is 3.5 MPa higher than atmospheric pressure. The pressure difference causes the gaseous precursor to flow and diffuse around and inside the turbine rotor fiber preform, prompting the precursor gas to undergo a chemical reaction in the pores, depositing to form a ceramic matrix, filling the pores and forming a dense composite material.

[0038] Specifically, the fiber preform is an integral turbine blade disk structure, that is, 33 turbine working blades and a turbine disk are an integral structure, and the blades and other complex surfaces are three-dimensionally woven.

[0039] Specifically, the reaction gas uses a gaseous precursor with a chemical ratio of Si:C of 1:1 as the deposited SiC, the gaseous precursor is composed of trichloromethylsilane, H2 is used as a carrier gas, and is mixed into the reaction gas at a ratio of H2 / trichloromethylsilane of 4 to 8.

[0040] Embodiment three;

[0041] See also Figures 1 to 3 , an aero-engine ceramic-based turbine rotor and a preparation process thereof of the present invention, this embodiment also includes;

[0042] The steps include:

[0043] S1: Fiber preform: Ordinary 2D plain woven fiber cloth is stacked from bottom to top, and the angle between the upper and lower layers of fiber is fixed at 35°. After the layers reach the required thickness, the fiber layers are sutured and connected by puncture to complete the formation of the turbine rotor preform;

[0044] S2: Ceramic preparation: The turbine rotor uses SiC substrate, and the turbine rotor preform is placed in a reaction chamber at 1100°C and 8KPa;

[0045] S3: mixing reaction gas and carrier gas into the reaction chamber;

[0046] S4: The pressure above the preform in the reaction chamber is atmospheric pressure, and the precursor pressure is 3.5 MPa higher than atmospheric pressure. The pressure difference causes the gaseous precursor to flow and diffuse around and inside the turbine rotor fiber preform, prompting the precursor gas to undergo a chemical reaction in the pores, depositing to form a ceramic matrix, filling the pores and forming a dense composite material.

[0047] Specifically, the fiber preform is an integral turbine blade disk structure, that is, 33 turbine working blades and a turbine disk are an integral structure, and the blades and other complex surfaces are three-dimensionally woven.

[0048] Specifically, the reaction gas uses a gaseous precursor with a chemical ratio of Si:C of 1:1 as the deposited SiC, the gaseous precursor is composed of trichloromethylsilane, H2 is used as a carrier gas, and is mixed into the reaction gas at a ratio of H2 / trichloromethylsilane of 4 to 8.

[0049] The above shows and describes the basic principle, main features and advantages of the present invention, and the standard parts used in the present invention can be purchased from the market, and special-shaped parts can be customized according to the description in the specification and the drawings. The specific connection methods of each part adopt the conventional means such as mature bolts, rivets, welding, etc. in the prior art. The machinery, parts and equipment all adopt the conventional models in the prior art, and the circuit connection adopts the conventional connection method in the prior art, which will not be described in detail here.

[0050] The above description of the disclosed embodiments enables one skilled in the art to implement or use the present invention. Various modifications to these embodiments will be apparent to one skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the present invention. Therefore, the present invention will not be limited to the embodiments shown herein, but rather to the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A process for preparing a ceramic-based turbine rotor for an aero-engine, comprising the following steps; S1: Fiber preform: Ordinary 2D plain fiber cloth is stacked from bottom to top, and the angle between the upper and lower layers of fiber is fixed at 30° to 45°. After the layers reach the required thickness, the fibers of each layer are sutured and connected by puncture to complete the formation of the turbine rotor preform; S2: The turbine rotor uses a SiC substrate, and the turbine rotor preform is placed in a reaction chamber at 900-1100°C and 4-10KPa; S3: mixing reaction gas and carrier gas into the reaction chamber; S4: The pressure above the preform in the reaction chamber is atmospheric pressure, and the precursor pressure is 3.5 MPa higher than atmospheric pressure.

2. The process for preparing a ceramic-based turbine rotor for an aero-engine according to claim 1, characterized in that: The fiber preform is an integral turbine blade disk structure.

3. The process for preparing a ceramic-based turbine rotor for an aero-engine according to claim 1, characterized in that: The reaction gas uses a gaseous precursor with a chemical ratio of Si:C of 1:1 as deposited SiC. The gaseous precursor consists of trichloromethylsilane, and H2 is used as a carrier gas, and is mixed into the reaction gas at a ratio of H2 / trichloromethylsilane of 4 to 8.

4. A basic turbine rotor, characterized in that: The turbine rotor is prepared by using the preparation process for a ceramic-based turbine rotor for an aero-engine as described in any one of claims 1 to 3.

5. An engine, characterized in that: The engine uses the ceramic-based turbine rotor of claim 4.

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