Aero-engine ceramic matrix composite fixed guider blade structure and forming method thereof

An aero-engine and composite material technology, which is applied in the direction of engine components, machines/engines, air transportation, etc., can solve the problems of unreasonable design of CMC guide blade prefabrication, poor blade root strength, etc., so as to reduce the risk of preparation and reduce the cost , good technical effect

Active Publication Date: 2020-05-05
XIAN XINGUI CERAMIC COMPOSITE MATERIAL CO LTD
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  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

[0006] In order to overcome the structural defects of unreasonable design of the existing CMC guide vane prefabricated body and poor blade

Method used

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  • Aero-engine ceramic matrix composite fixed guider blade structure and forming method thereof
  • Aero-engine ceramic matrix composite fixed guider blade structure and forming method thereof
  • Aero-engine ceramic matrix composite fixed guider blade structure and forming method thereof

Examples

Experimental program
Comparison scheme
Effect test

Example Embodiment

[0054] Example one

[0055] The dimensions of the CMC fixed guide vane in this embodiment are: 60 mm in length, 50 mm in width, and 120 mm in height. In this embodiment, the raw material of the ceramic matrix composite material is SiC fiber, and the raw materials used in the preparation process are trichloromethylsilane and H 2 , Ar gas, etc.

[0056] Specific steps are as follows:

[0057] (1) Preparation of preforms: SiC fibers are used to weave 2D plain weave cloth, and other types of preforms such as 2.5D and 3D can also be used. According to the size of the parts, the plain weave cloth is cut into a suitable size, and multiple layers of plain weave cloth are stacked and punctured in the stacking direction of the plain weave cloth. The puncture fiber uses the same SiC fiber to form a SiC fiber preform. The SiC fiber preform is fixed and formed using a mold.

[0058] (2) Preparation of the interface layer: the preform of step (1) is placed in a chemical vapor deposition furnace t...

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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 a forming method thereof, and belongs tothe technical field of aero-engine fixed guider preparation. According to the aero-engine ceramic matrix composite fixed guider blade structure and the forming method thereof, a turbine guider blade is prepared in an integrated assembly mode, parts are integrally assembled in a riveting mode, an assembly structure adopts a SiC ceramic matrix to achieve welding, and therefore the defect that the strength of a blade root is insufficient due to a conventionally sewed and formed prefabricated body can be overcome; protrusions on a first lower edge plate and a second lower edge plate are directly riveted with the blade in a limiting mode and then riveted with a third lower edge plate and a fourth lower edge plate through riveting rivets, a SiC ceramic matrix is deposited on a whole product through a CVI technology, and therefore integrated preparation is completed; and in this way, the SiC/SiC high-pressure guide blade has good manufacturability, a margin plate and the blade are simpler toprepare 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 [0001] The invention relates to a structure and molding of a fixed guide blade, in particular to an assembly structure and a forming method of a fixed guide blade of a ceramic matrix composite material for an aero-engine, and belongs to the technical field of preparation of an aero-engine fixed guide. Background technique [0002] The demand for high-performance aeroengines for military and civilian aircraft is increasingly urgent. There are two main ways to improve the performance of the engine. One is to increase the boost ratio of the compressor, and the other is to increase the gas temperature at the turbine inlet. As the gas temperature at the turbine inlet increases, the high temperature components of the engine need to bear greater heat load. The gas temperature in front of the turbine of an engine with a thrust-to-weight ratio of 10 in developed countries can reach 1850-1950K, which can generate more than double the thrust of the previous generation o...

Claims

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Application Information

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IPC IPC(8): F01D9/04F01D25/00C04B35/565C04B35/622C04B35/80
CPCF01D9/041F01D25/005F01D25/00C04B35/565C04B35/622Y02T50/60
Inventor 涂建勇王佳民许建锋刘梦珠郭洪强成来飞
Owner XIAN XINGUI CERAMIC COMPOSITE MATERIAL CO LTD
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