GB2441865A

GB2441865A - Hybrid ceramic matrix composite turbine vane assembly

Info

Publication number: GB2441865A
Application number: GB0717266A
Authority: GB
Inventors: Jeffrey H Boy; Dennis J Landini
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2006-09-14
Filing date: 2007-09-05
Publication date: 2008-03-19
Also published as: GB0717266D0; DE102007039402A1; JP2008069782A; CA2599626A1

Abstract

A hybrid turbine blade 10 includes a composite core 16 enclosed within a composite shell 14. The composite core includes a mixture of silicon carbide, silicon and carbon; the composite shell includes a wound or woven ceramic silicon carbide fiber material. A method of forming the blade may include the steps of (a) shaping a carbon preform; (b) reacting the preform in a furnace with silicon at a temperature above the melting point of silicon to form a CMC composite core component 16 made up of a mixture of silicon carbide, silicon and carbon; (c) locating a temporary mandrel core component 18 adjacent the CMC composite core component 16 to form a multi-core assembly; (d) forming a composite shell 14 by laying up prepreg tapes around the multi-core assembly; and (e) removing the temporary mandrel core 18. The shell 14 may then be subjected to melt infiltration of molten silicon. Alternatively, the shell and the core may be melt-infiltrated simultaneously.

Description

2441865

HYBRID CERAMIC MATRIX COMPOSITE TURBINE VANE ASSEMBLY

AND RELATED METHOD

This invention relates generally to turbine technology and more specifically, to high temperature turbine vane assemblies.

Ceramic matrix composites (CMCs) offer the potential for higher operating temperatures than comparable metal alloy materials. CMCs, however, are not as strong as metals and thus require thicker cross sections to compensate. The use of CMCs for various turbine components is known and is reflected in the patent literature. For example, U.S. Patent No. 6,709,230 discloses a hybrid composite vane for a gas turbine engine having a CMC airfoil member bonded to a substantially solid core member with cooling channels.

U.S. Patent No. 6,670,021 discloses a monolithic ceramic attachment bushing incorporated into a CMC component. The component in one embodiment may include an inner core formed as an attachment bushing or a wear pad made of either silicon carbide or silicon nitride monolithic ceramic that is embedded within a CMC shell.

U.S. Patent No. 6,607,358 discloses a multi-component hybrid turbine blade where the airfoil portion comprises a composite section having a first density and an insert section having a second density less than the first density.

U.S. Patent No. 6,451,416 discloses a hybrid monolithic ceramic and ceramic matrix composite airfoil where the airfoil comprises a temperature resistant, monolithic ceramic exterior layer with a tough, high impact resistant, fiber reinforced ceramic matrix composite interior layer.

In the manufacture of composite turbine vanes, it has been the practice to form the airfoil about a solid carbon core, by laying up prepreg tapes about the solid carbon

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core, with additional tapes in, for example, the trailing edge region of the vane, such that this area of the vane was made up exclusively of the prepreg tapes. Because of the relatively high cost of this process, it is desirable to develop a less costly process that does not impact on vane performance.

In accordance with one exemplary but non-limiting embodiment of this invention, a hybrid turbine vane or blade assembly comprises a composite multi-core assembly in combination with a different composite shell assembly. The multi-core assembly includes a first solid composite core component fabricated from a short-chopped carbon preform, machined to shape, that undergoes conversion to SiC by reactive melt infiltration with molten silicon, yielding a composite consisting of a mixture of silicon carbide, silicon and carbon. This first core component is mated with a second but temporary solid mandrel core component, which may be carbon, aluminum or some other material, to form the blade multi-core assembly. The shell component includes a woven or wound ceramic matrix composite material wrapped about the multi-core assembly.

The shell is prepared by a conventional prepreg process and shaped to form the vane or blade by laying up the prepreg tapes around the multi-core assembly. An example of the process is described in U.S. Patent Nos. 6,024,898 and 6,258,737. After forming the shell, the temporary mandrel core component is removed, and thereafter, the remaining composite core component and the wound or woven prepreg shell are subjected to melt infiltration. The melt infiltration densifies the outer prepreg shell by filling it's porosity. The molten silicon also softens the silicon in the composite core section forming an integral bond with the outer shell.

In an alternative example, a green carbon preform (not yet reacted with silicon) is used in the multi-core assembly along with the temporary mandrel core. After the prepreg tapes are wound about the core assembly and the temporary mandrel core is removed, the remaining composite core and shell are melt-infiltrated together. The melt infiltration densifies the outer shell by filling the porosity. The infiltration of the

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"green" carbon preform by molten silicon leads to the chemical reaction between the carbon and silicon to form silicon carbide insitu. This conversion of the carbon preform forms a mixture of silicon carbide, carbon and silicon. The molten silicon in the reacted preform and the outer prepreg shell forms an integral bond between the reacted core to the prepreg shell. The internal space created by removal of the temporary mandrel core can be utilized for cooling the blade.

By forming a particular portion of the vane, e.g., the trailing edge region, with a CMC composite core component, the more costly process of laying up prepreg tapes to form this portion of the vane is substantially eliminated in that the tapes can now be wound about a multi-core assembly, with one part of the multi-core assembly remaining in place in the trailing edge region and bonded to the shell.

Accordingly, in one aspect, the invention relates to a hybrid turbine blade comprising a composite core enclosed within a composite shell, the composite core including a mixture of silicon carbide, silicon and carbon; the composite shell including a wound or woven ceramic carbide fiber material.

In another aspect, the invention relates to a method of forming a turbine blade comprising: (a) shaping a carbon preform and reacting the preform in a furnace with silicon at a temperature above the melting point of silicon to thereby form a CMC composite core consisting of a mixture of silicon carbide, silicon and carbon; (b) forming a composite shell by laying up prepeg tapes around the CMC composite core and a discrete mandrel core; and (c) removing the discrete mandrel core.

In still another aspect, the invention relates to a hybrid turbine blade comprising a CMC material core enclosed within a composite shell, the composite core including a mixture of silicon carbide, silicon and carbon; the composite shell including a wound or woven ceramic carbide fiber material.

The invention will now be described in connection with the drawing figures identified below, in which:

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FIGURE 1 is a schematic cross section illustrating the core and shell components of a turbine vane in accordance with an exemplary embodiment of the invention; and

FIGURE 2 is a micrograph showing the melt infiltrated core component of Figure 1.

With reference to Figure 1, a hybrid turbine vane or blade assembly 10 in accordance with an exemplary but non-limiting embodiment of the invention includes a composite core component 16 and composite shell 14. The core component 16 may comprise a known solid composite fabricated from a short-chopped carbon preform. A "green" carbon preform in this exemplary embodiment is initially machined to shape and the core component is then reacted in a vacuum furnace with sufficient silicon (greater than 2.7 times the weight of the carbon preform), at a temperature above the melting point of silicon to react the carbon to form a CMC composite core component 16 that consists of a mixture of silicon carbide, silicon and carbon. One such composite suitable for use in this process is marketed under the trade name Cesic®. Note that the Cesic® core component 16 does not fill the entire interior of the vane or blade, but only in areas where needed for structural support, for example, in the trailing edge region of the airfoil. During manufacture, however, and as explained further below, the Cesic® core component is utilized in combination with a temporary solid mandrel core component 18 (thus forming a multi-core assembly) that is later removed.

The shell 14 is comprised of a wound or woven ceramic silicon carbide fiber prepared in accordance with a conventional prepreg process and shaped in the form of a vane airfoil by laying up the prepreg tapes about the multi-core assembly that includes the CMC composite Cesic® core component 16 and the adjacent, discrete temporary mandrel core component 18. The discrete, temporary mandrel core 18 fills the remaining portion of the vane interior, but is subsequently removed after the shell 14 is wound about the multi-core assembly. Because the shell is open-ended, the temporary mandrel core component 18 may be removed simply by sliding it out of the shell 14 from either of the open ends. The CMC Cesic® core component 16 and

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prepreg shell 14 are then subjected to melt infiltration of molten silicon. The melt infiltration densifies the outer prepreg shell 14 by filling it's porosity. The molten silicon also softens the silicon in the CMC Cesic® core component 16 forming an integral bond with the outer prepreg shell 14.

In an alternative example, the reaction of a green preform core 16 into a composite mixture of silicon carbide, silicon and carbon, and the melt infiltration of the shell 14 may be conducted simultaneously, in the same vacuum furnace cycle with molten silicon for the densification and bonding of the shell 14 to the core 16.

After the prepreg tapes are wound about the core assembly and the temporary mandrel core component 18 is removed, the remaining core component 16 and shell 14 are melt-infiltrated together. The melt infiltration densifies the outer shell 14 by filling the porosity. The infiltration of the "green" carbon by molten silicon leads to the chemical reaction between the carbon and silicon to form silicon carbide insitu. This conversion of the carbon preform forms a mixture of silicon carbide, carbon and silicon. The molten silicon in the reacted preform and the outer prepreg shell forms an integral bond between the reacted core component 16 and the prepreg shell 14. The internal space created by removal of the temporary mandrel core can be utilized use for cooling the blade

Figure 2 is a photomicrograph illustrating the melt-infiltrated CMC composite core component 16.

While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.

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Claims

CLAIMS:

1. A hybrid turbine blade comprising a CMC material core enclosed within a composite shell, said composite core including a mixture of silicon carbide, silicon and carbon; said composite shell including a wound or woven ceramic carbide fiber material bonded to the core.

2. The blade of claim 1 wherein said CMC material core is formed from a chopped-carbon preform reacted with silicon.

3. The blade of claim 1 or claim 2 wherein said CMC material core and said shell are melt infiltrated with molten silicon to bond the shell to the core.

4. The blade of any preceding claim wherein said CMC material composite core occupies only a portion of an interior space of the blade defined by said shell.

5. The blade of claim 4 wherein said CMC material core occupies a trailing edge region of said blade.

6. A method of forming a turbine blade comprising:

(a) shaping a carbon preform;

(b) reacting the preform in a furnace with silicon at a temperature above the melting point of silicon to thereby form a CMC composite core component made up of a mixture of silicon carbide, silicon and carbon;

(c) locating a temporary mandrel core component adjacent said CMC composite core component to form a multi-core assembly;

(d) forming a composite shell by laying up prepreg tapes around said multi-core assembly; and

(e) removing said temporary mandrel core.

7. The method of claim 6 and further comprising:

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(f) subjecting said shell to melt infiltration of molten silicon to densify said shell and bond said composite core to said shell.

8. The method of claim 6 or claim 7 wherein said furnace comprises a vacuum furnace.

9. A method of forming a turbine blade comprising:

(a) shaping a first CMC preform core component from chopped carbon;

(b) combining said first CMC core component with a second temporary mandrel core component to form a multi-core assembly;

(c) wrapping said multi-core assembly with prepreg tapes to form an outer shell;

(d) removing said temporary mandrel core component; and

(e) simultaneously melt-infiltrating said outer shell and said first CMC core component with silicon to chemically react said first CMC core component and to densify said outer shell and bond said first CMC core component to said outer shell.

10. The method of claim 9 wherein step (d) is carried out at a temperature above the melting point of silicon in a vacuum furnace.

11. A hybrid turbine blade substantially as hereinbefore described with reference to the accompanying drawings.

12. A method of forming a turbine blade substantially as hereinbefore described with reference to the accompanying drawings.

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