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US-12617730-B2 - Direct bonded environmental barrier coatings for sic/sic composites and methods for preparing the same

US12617730B2US 12617730 B2US12617730 B2US 12617730B2US-12617730-B2

Abstract

A method of preparing a ceramic matrix composite (CMC) article is disclosed. The method includes depositing a first layer of a coating composition directly onto a surface of a silicon carbide fiber-reinforced silicon carbide matrix (SiC/SiC) composite substrate, with the coating composition comprising a rare earth silicate and a sintering aid. The method also includes heating the first layer to sinter the coating composition to form an environmental barrier coating (EBC) adjacent the SiC/SiC composite and a transition layer integrally bonded to and between the substrate and the EBC. CMC articles prepared according to the method, including coated turbomachine components, are also disclosed.

Inventors

  • Mehrad Mehr
  • Bahram Jadidian

Assignees

  • HONEYWELL INTERNATIONAL INC.

Dates

Publication Date
20260505
Application Date
20220330

Claims (17)

  1. 1 . A method of preparing a ceramic matrix composite (CMC) article, the method comprising: depositing a first layer of a coating composition directly onto a surface of a silicon carbide fiber-reinforced silicon carbide matrix (SiC/SiC) composite substrate, the coating composition comprising a rare earth silicate and a sintering aid; and heating the first layer to sinter the coating composition to form an environmental barrier coating (EBC) adjacent the SiC/SiC composite substrate and a transition layer integrally bonded to and between the SiC/SiC composite substrate and the EBC, thereby preparing the CMC article, wherein the transition layer comprises a reaction product of the rare earth silicate of the coating composition and the SiC/SiC composite substrate, wherein the transition layer has a thickness of from 3 to 10 mils, and the EBC has a thickness of from 1.5 to 10 mils.
  2. 2 . The method of claim 1 , wherein the surface of the SiC/SiC composite substrate on which the first layer of the coating composition is deposited is free from a bond coating.
  3. 3 . The method of claim 1 , wherein the first layer of the coating composition is deposited via direct application to the SiC/SiC composite substrate by doctor blading or airbrushing.
  4. 4 . The method of claim 3 , wherein the coating composition is in the form of a slurry or paste, and wherein the method further comprises preparing the slurry or paste prior to depositing the first layer thereof by combining particles of the rare earth silicate, particles of the sintering aid, a carrier vehicle, and optionally a binder.
  5. 5 . The method of claim 1 , wherein heating the first layer of the coating composition comprises: a burnout phase in which the first layer of the coating composition is heated to a first temperature (T 1 ) within a first time period, and subsequently a sintering phase in which the first layer of the coating composition is heated to a second temperature (T 2 ) higher than T 1 within a second time period.
  6. 6 . The method of claim 5 , wherein during the sintering phase, the first layer of the coating composition is substantially free from: (i) carbon; (ii) free silicon; or (iii) both (i) and (ii).
  7. 7 . The method of claim 5 , wherein the second temperature (T 2 ) and the second time period are selected to prepare the transition layer comprising the reaction product of the rare earth silicate of the coating composition and the SiC/SiC composite substrate.
  8. 8 . The method of claim 1 , wherein the rare earth silicate is further defined as a rare earth disilicate (RE 2 Si 2 O 7 ), and wherein the first coating composition comprises from about 80 to about 99 wt. % of the rare earth disilicate, based on the total weight solids of the first coating composition.
  9. 9 . The method of claim 8 , wherein the rare earth silicate comprises ytterbium disilicate (Yb 2 Si 2 O 7 ).
  10. 10 . The method of claim 9 , wherein the coating composition comprises from about 1 to about 20 wt. % of the sintering aid, based on the total weight solids of the first coating composition.
  11. 11 . The method of claim 1 , wherein the sintering aid comprises alumina (Al 2 O 3 ).
  12. 12 . The method of claim 1 , wherein the sintering aid is substantially free from: (i) carbon; (ii) silicon; or (iii) both (i) and (ii).
  13. 13 . The method of claim 1 , further comprising depositing a layer of a second coating composition over the EBC; and heating the second layer to form a second EBC bonded to the EBC prepared from the first coating composition.
  14. 14 . The method of claim 13 , wherein the first and second coating compositions are substantially the same.
  15. 15 . The method of claim 1 , wherein the SiC/SiC composite substrate is a part body, and wherein depositing the first layer of the coating composition comprises encapsulating at least a portion of the part body with the coating composition.
  16. 16 . The method of claim 1 , further comprising disposing a thermal barrier coating (TBC) over the EBC.
  17. 17 . The method of claim 1 , wherein: the EBC has a first coefficient of thermal expansion (CTE); the transition layer has a second CTE; and the SiC/SiC composite substrate has a third CTE, wherein the second CTE more closely matches the third CTE than the first CTE matches the third CTE, and this closer match is achieved because the transition layer includes a reaction product formed from the rare earth silicate of the coating composition and the SiC/SiC composite substrate.

Description

TECHNICAL FIELD The present disclosure relates generally to protective coatings for engine components and, more specifically, to direct bonded environmental barrier coatings (EBC) for silicon-based ceramic substrates using an integrally formed transition layer, parts prepared therewith, methods of preparing such coated parts. BACKGROUND Gas turbine engines are used as the primary power source for various kinds of aircraft and other vehicles. The engines may also serve as auxiliary power sources that drive air compressors, hydraulic pumps, and industrial electrical power generators. Demands for engine efficiency have led to higher operating temperatures and new component materials suitable for use under resulting conditions. For example, parts fabricated from ceramic matrix composites (CMCs) are increasingly employed, as such materials exhibit improved mechanical, physical, and chemical properties at high temperatures compared to parts made from conventional metal alloys (e.g. nickel-based superalloys). CMCs are composed of a ceramic matrix phase and reinforcing fibers or particles, and Common materials used for both the matrix and reinforcing materials include Carbon (C), silicon carbide (SiC), alumina (Al2O3) and mullite (Al2O3—SiO2). As such, CMCs are typically named using a combination of the type of fiber/type of matrix. For example, C/C refers to carbon fiber-reinforced carbon, C/SiC for carbon fiber-reinforced silicon carbide, and SiC/SiC for silicon carbide fiber-reinforced silicon carbide. Compared to nickel-based superalloys, CMC materials can be prepared with extremely high thermal, mechanical, and chemical stability, while also maintaining a high strength-to-weight ratio. As such, CMC components, such as turbine blades, shrouds, and nozzles, may be employed to allow gas turbine engines to operate more efficiently and at higher temperatures. Unfortunately, CMCs do not exhibit acceptable environmental durability in typical combustion environments. In particular, water vapor produces during the combustion process can react with the CMC, e.g. to form gaseous reaction products such as Si(OH)4, leading to loss of material, erosion, and/or surface recession of the CMC. To overcome these limitations, environmental barrier coatings (EBCs) are used on CMC components used in turbine engine hot zones, such as high-pressure turbine (HPT) shrouds and nozzles, to protect the base ceramic from the steam in the gas flow. Such components may also include a thermal barrier coating (TBC) (i.e., separate from any EBC), to insulate and minimize thermal impact on the engine structures due to temperature cycling, exposure to airborne contaminants such as calcia-mangesia-alumina-silicate (CMAS), etc. Typically, formation of EBC is carried out after machining the structure of the component to be coated to a desired shape. For example, silicon-based ceramic substrates are generally formed through sintering processes, in which silicon-based powder and a sintering aid are shaped in a mold by batch powder addition until a desired thickness is reached. The powder is typically uniform in composition during the process, with relatively low amounts of sintering aid (e.g. less than about 5 wt. %, based on the total weight of the powder composition). The powder then cold pressed in the mold is (i.e., green body formation), and then fused (e.g. via glass encapsulation and subsequent sintering, pressureless sintering, or polymer infiltration and pyrolysis). Once sintered, the substrate is machined and annealed to achieve a target shape and tolerance parameters for the particular part being prepared. EBCs must be formed after sintering and machining of the substrate in order to both maintain dimensional tolerances in the part, as well as to prevent compromising the coverage of substrate by the EBC via machining the coating. Additionally, EBCs are typically formed via processes incompatible with the substrate sintering, such as plasma spray and electron beam physical vapor deposition (EB-PVD). Unfortunately, conventional EBCs have practical limitations preventing use in desirable applications. For example, HPT surface temperatures may be targeted in excess of 2600° F. to improve performance characteristics and engine efficiency. However, conventional EBCs based on rare earth disilicates have a temperature limit of around 2400° F. in the HPT environment, above which the disilicate decomposes to a more stable monosilicate phase that is porous and has a higher coefficient of thermal expansion (CTE) than the substrate or the EBC itself. The porosity of the resulting monosilicate layer continuously exposes the underlying EBC to exhaust gasses, perpetuating the growth of the layer during operation. Moreover, the CTE differential may lead to stresses at the coating interface during cyclic heating to operating temperatures, which can result in delamination of the coating. Compounding the performance issues above, conventional EBC materials oft