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Structure and Forming of Aeroengine Ceramic Matrix Composite Fixed Guide Vane

A technology of aero-engine and composite materials, applied in engine components, machines/engines, stators, etc., can solve problems such as poor blade root strength, unreasonable design of CMC guide blade prefabricated bodies, etc., to reduce costs, reduce manufacturing risks, The effect of easy guarantee of dimensional accuracy

Active Publication Date: 2021-12-14
XIAN XINGUI CERAMIC COMPOSITE MATERIAL CO LTD
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  • Abstract
  • Description
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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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  • Structure and Forming of Aeroengine Ceramic Matrix Composite Fixed Guide Vane
  • Structure and Forming of Aeroengine Ceramic Matrix Composite Fixed Guide Vane
  • Structure and Forming of Aeroengine Ceramic Matrix Composite Fixed Guide Vane

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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 fixed guides of aero-engines, and relates to a blade structure of a fixed guide made of a ceramic matrix composite material of an aero-engine and its forming. To overcome the structural defects of unreasonable design of the existing CMC guide vane prefabrication and poor blade root strength, the turbine guide vane including multi-stage guide vanes is prepared by integrated assembly method, and the parts are integrated and assembled by riveting, and the assembly structure is realized by SiC ceramic matrix "Welding"; and make full use of the effective space between the upper edge plate and the outer casing, install and position the blade body and the upper edge plate through limit rivets and limit platforms, and use CVI technology to deposit SiC ceramic substrates on the entire product , to complete the integrated preparation. This structure makes the complex SiC / SiC multi-connected high-pressure guide vane have good manufacturability, the preparation of the edge plate and the vane is simpler, the installation is more reliable, and the dimensional accuracy of the final product is easier to guarantee. This structure can be fully applied to medium and large engines and fifth-generation aircraft.

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 Patents(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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