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Ceramic matrix composite elastic sealing element for aero-engine and preparation method

A technology of elastic seals and aero-engines, applied in the direction of engine seals, engine components, mechanical equipment, etc., can solve the problems of easy failure, high density, poor temperature resistance, etc., and achieve the effect of highly reliable sealing connection

Inactive Publication Date: 2020-08-11
NORTHWESTERN POLYTECHNICAL UNIV
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

[0011] In order to avoid the deficiencies of the prior art, the present invention proposes a ceramic-matrix composite elastic seal for aero-engines and a preparation method thereof, which solves the defects of conventional high-temperature elastic seals such as high density, poor temperature resistance, and easy failure. To meet the development needs of high-temperature sealing of hot-end components of high-performance aero-engines in the future

Method used

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  • Ceramic matrix composite elastic sealing element for aero-engine and preparation method

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Experimental program
Comparison scheme
Effect test

Embodiment 1

[0029] Step 1. Using C fiber to prepare an elastomer preform, the structure of the preform is 2D, and the thickness of the preform is 4 mm.

[0030] Step 2. Place the prefabricated body between the graphite inner mold and the outer mold and press it tightly. The graphite mold surface is punched with a hole diameter of 2 mm.

[0031] Step 3. Place the fiber prefabricated body with the mold in the CVD interface deposition equipment, the interface layer deposition temperature is 500-1100°C, the vacuum of the deposition furnace is 3-50kPa, 90L / min Ar gas is used as the protective gas, trichloride Boron is used as the precursor gas of the interface layer, the gas flow rate of boron trichloride is 140L / min, the deposition time is 30h, and the interface thickness is controlled at 600μm;

[0032] Step 4. Place the preform after the deposition interface in the CVI deposition equipment, and start SiC matrix deposition. The SiC ceramic substrate deposition temperature is 1100-1400°C, t...

Embodiment 2

[0038] Step 1. SiC fibers are used to prepare an elastomer preform, the structure of which is 2.5D, and the thickness of the preform is 5mm.

[0039] Step 2. Place the prefabricated body between the graphite inner mold and the outer mold and press it tightly. The surface of the graphite mold is perforated with a diameter of 4mm.

[0040] Step 3. Place the fiber preform with the mold in the CVD interface deposition equipment. The interface layer deposition temperature is 500-1100°C, the vacuum of the deposition furnace is 3-50kPa, 180L / min Ar gas is used as the protective gas, trichloride Boron is used as the interface layer precursor gas, the gas flow rate of boron trichloride is 300L / min, the deposition time is 40h, and the interface thickness is controlled at 650μm;

[0041]Step 4. Place the preform after the deposition interface in the CVI deposition equipment, and start SiC substrate deposition. SiC ceramic substrate deposition temperature is 1100-1400°C, the deposition ...

Embodiment 3

[0047] Step 1. The elastomer preform is prepared by using fibers such as boron nitride. The preform has a 3D structure and the thickness of the preform is 6mm.

[0048] Step 2. Place the prefabricated body between the graphite inner mold and the outer mold and press it tightly. The surface of the graphite mold is perforated with a diameter of 5 mm.

[0049] Step 3. Place the fiber preform with the mold in the CVD interface deposition equipment. The interface layer deposition temperature is 500-1100°C, the vacuum of the deposition furnace is 3-50kPa, 200L / min Ar gas is used as the protective gas, trichloride Boron is used as the interface layer precursor gas, the gas flow rate of boron trichloride is 450L / min, the deposition time is 50h, and the interface thickness is controlled at 700μm;

[0050] Step 4. Place the preform after the deposition interface in the CVI deposition equipment, and start SiC substrate deposition. The SiC ceramic substrate deposition temperature is 110...

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Abstract

The invention relates to a ceramic matrix composite elastic sealing element for an aero-engine and a preparation method. The sealing element is a revolving body or a long strip, the cross section of the sealing element is omega-shaped, and the sealing element is riveted to the surface of a hot end part of an aero-engine by adopting a ceramic matrix composite rivet. The upper edge of the omega is in close contact with the other sealing surface. The opening size L of the sealing element can be compressed or bounced along with the rising and falling of the environment temperature of the engine and the change of environment conditions such as airflow impact and mechanical vibration of the engine, so that the nested structure of the hot end part or the effective sealing of the junction surfaceis realized. The density of the prepared ceramic matrix composite elastic sealing element is larger than or equal to 2.4g / cm<3>, and the porosity of the material is smaller than or equal to 6%. The strength retention rate is greater than or equal to 95% under the temperature condition of 1650K, and the cracking stress of the sealing material matrix is greater than or equal to 120MPa; when the bearing stress is smaller than or equal to 100MPa, the linear resilience characteristic is kept; and the service life is more than or equal to 5000 hours at the temperature of 1650K.

Description

technical field [0001] The invention belongs to the technical field of mechanical high-temperature sealing, and relates to a ceramic-matrix composite material elastic sealing member for an aero-engine and a preparation method thereof. Background technique [0002] With the development of high-performance aircraft technology, there is an urgent demand for high-performance engines in the aerospace field. Studies have shown that increasing the turbine inlet gas temperature has become an effective way to improve engine performance, and all countries have invested heavily in exploration, and have successively carried out relevant technical assessment and verification. At present, the gas temperature before the turbine of an engine with a thrust-to-weight ratio of 10 in developed countries has reached 1850-1950K. In the future, the thrust-to-weight ratio of the fifth-generation aeroengine will be about 15-20, and the gas temperature before the turbine will reach 2200-2400K; techno...

Claims

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Application Information

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Patent Type & Authority Applications(China)
IPC IPC(8): C04B35/573C04B35/577C04B35/80C04B35/622F16J15/30
CPCC04B35/573C04B35/622C04B2235/524C04B2235/5244C04B2235/5248C04B2235/5252C04B2235/5256C04B2235/614C04B2235/77C04B2235/96F16J15/30
Inventor 董宁刘小冲涂建勇何江怡刘持栋孙肖坤成来飞
Owner NORTHWESTERN POLYTECHNICAL UNIV