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Aero-engine ceramic matrix composite fixation guider blade structure and molding method thereof

An aero-engine and composite material technology, applied in the direction of engine components, machines/engines, stators, etc., can solve problems such as poor blade root strength, unreasonable design of CMC guide blade prefabrication, etc., to reduce costs, reduce manufacturing risks, The effect of improving component quality

Active Publication Date: 2020-04-07
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 root strength, the present invention provides a novel ceramic matrix composite material fixed guide vane structure and its forming method

Method used

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  • Aero-engine ceramic matrix composite fixation guider blade structure and molding method thereof
  • Aero-engine ceramic matrix composite fixation guider blade structure and molding method thereof
  • Aero-engine ceramic matrix composite fixation guider blade structure and molding method thereof

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Embodiment Construction

[0054] The present invention will be described in detail below in conjunction with the accompanying drawings and specific embodiments. The specific embodiments of the present invention further illustrate the thinking, technical problems to be solved, features and effects of the technical solutions of the present invention. The description of the embodiments does not constitute a limitation of the present invention. In addition, the technical features involved in the embodiments of the present invention may be combined with each other as long as they do not conflict with each other.

[0055] The blade structure of the aeroengine ceramic-based composite material fixed guide in this embodiment is mainly aimed at triple guide vanes, including the blade body 1, the upper edge plate assembly 2, the limit rivet 4, the riveting rivet 5 and the limit platform 6.

[0056] Wherein the airfoil 1 includes the airfoil body and the limiting plate 7 arranged on the upper end surface of the a...

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Abstract

The invention belongs to the technical field of preparation of an aero-engine fixation guider, relates to an aero-engine ceramic matrix composite fixation guider blade structure and a molding method thereof, and is to overcome the structural defects that an existing CMC guider blade prefabricated body is unreasonable in design and poor in blade root strength. A turbine guider blade manufactured in an integrated assembly mode comprises multiple stages of guide blades, parts are integrally assembled in a riveting mode, and an assembly structure adopts a SiC ceramic matrix to achieve welding; aneffective space between an upper edge plate and an outer casing is fully utilized, the blade body and the upper edge plate are installed and positioned through limiting rivets and a limiting table, the SiC ceramic matrix is deposited on the whole product through the CVI technology, and integrated preparation is completed; by means of the structure, the complex SiC / SiC multi-connected high-pressure guide blade has good manufacturability, the edge plate and the blade are easier to manufacture and more reliable to install, and the dimensional precision of a final product is easier to guarantee;and the structure can be fully applied to medium-sized and large-sized engines and five-generation 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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Patent Type & Authority Applications(China)
IPC IPC(8): F01D9/04F01D25/00C04B35/565C04B35/80C04B35/622
CPCF01D9/041F01D25/005C04B35/565C04B35/62868C04B35/622C04B2235/5244C04B2235/5256C04B2235/608F05D2230/314F05D2300/6033
Inventor 涂建勇王佳民陈旭何江怡王文红卜石成来飞
Owner XIAN XINGUI CERAMIC COMPOSITE MATERIAL CO LTD
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