WO2020018090A1

WO2020018090A1 - Hybrid components having an intermediate ceramic fiber material

Info

Publication number: WO2020018090A1
Application number: PCT/US2018/042706
Authority: WO
Inventors: Jan H. MARSH; Evan C. LANDRUM; David J. Mitchell
Original assignee: Siemens Aktiengesellschaft; Siemens Energy, Inc.
Priority date: 2018-07-18
Filing date: 2018-07-18
Publication date: 2020-01-23

Abstract

The present invention relates to a turbine nozzle (18) that includes a first body (102) of a metal material (104) and a second body (106) of a ceramic matrix composite material (108). A gap (110) is disposed between the first body (102) and the second body (106) and a ceramic fiber material (112) is disposed within the gap (110). The ceramic fiber material (112) is effective to inhibit air flow into the gap (110) between the first body (102) and the second body (106), and transfer a thermal and mechanical load between the first body (102) and the second body (106).

Description

HYBRID COMPONENTS HAVING AN INTERMEDIATE CERAMIC FIBER MATERIAL

FIELD

The present invention relates to high temperature structures, and more particularly to a hybrid component comprising an intermediate ceramic fiber material between a ceramic matrix composite (CMC) portion and a metal portion of the component.

BACKGROUND

Gas turbines comprise a casing or cylinder for housing a compressor section, a combustion section, and a turbine section. A supply of air is compressed in the compressor section and directed into the combustion section. The compressed air enters the combustion inlet and is mixed with fuel. The air/fuel mixture is then combusted to produce a high temperature and pressure (working) gas. This working gas is then ejected past the combustor transition and injected into the turbine section of the turbine.

The turbine section comprises rows of vanes which direct the working gas to the airfoil portions of the turbine blades. The working gas travels through the turbine section, causing the turbine blades to rotate, thereby turning the rotor. The rotor is also attached to the compressor section, thereby turning the compressor and also an electrical generator for producing electricity. A high efficiency for the combustion turbine is achieved by heating the gas flowing through the combustion section to as high a temperature as is practical. The hot gas, however, may degrade the various metal turbine components, such as the combustor, transition ducts, vanes, ring segments and turbine blades that it passes when flowing through the turbine.

For this reason, strategies have been developed to protect hot gas path components from extreme temperatures, including the development and selection of high temperature materials adapted to withstand these extreme temperatures, and cooling strategies to keep the components adequately cooled during operation. For one, ceramic matrix composite (CMC) materials have been developed with high temperature resistance. CMC materials include a ceramic matrix reinforced with ceramic fibers. While CMC materials provide excellent thermal protection properties, the mechanical strength of CMC materials is still notably less than that of corresponding high temperature superalloy materials.

For this reason, proposed solutions have also added strengthening materials to the CMC material or supported the CMC material with a material having a greater mechanical strength. For example, as shown in FIG. 1 , some gas turbine components 60, e.g., vane 65, include a CMC body 62 having a metal body 64 extending radially therethrough for added mechanical strength. The metal body may comprise a spar 67 or platforms 72, 74 as illustrated. In such components and as shown in FIG. 2, directly interfacing the CMC body 62 and the metal body 64 allows for desirable heat and load transfer between the CMC body 62 and the metal body 64 during high temperature operation. Flowever, as desirable as it would be, the direct interface of the CMC body 62 and the metal body 64 is not possible in reality; there must be a gap between the two materials. First, the component surfaces will likely have some degree of imperfections or unevenness. Allowing for at least minimal manufacturing tolerances for each material requires that there must be some gap 66 between the CMC body 62 and the metal body 64 as shown in FIG. 3.

In addition, due to the thermal expansion differences between metal and CMC, a direct interface would render the CMC body 62 prone to cracking and failure as the metal body 64 expands to a greater degree than the CMC body 62 during high temperature operation and during thermal transitions as the components heat up and cool down. Accordingly, in this example, an even larger gap 68 is provided between the CMC body 62 and metal body 64 in order to accommodate thermal expansion of the metal body 64 toward the CMC body 62 as shown in FIG. 4. Flowever, the gap 68 reduces the thermal/mechanical load transfer between the CMC body 62 and the metal body 64, which may ultimately result in hot spots, high stress zones, and failure of the CMC material.

Further, in many components, such as in a vane 65 shown in FIGS. 1 and 5, the CMC body 62 must typically be secured to the corresponding metal body 64 to secure and properly support the CMC body 62. In this instance, the CMC body 62, e.g., a CMC shroud portion 70, is secured to a metal platform 72 and/or 74 (in the case of the vane 65) by a suitable structure, e.g., clamp 76 or the like. For the reasons set forth above (e.g., thermal expansion and manufacturing tolerances) plus the addition of the securing structure, clamp 76, there is a gap 68 disposed between the CMC shroud portion 70 and the metal platform 72 and 74. During normal operation of a gas turbine, hot gases may travel through the gap 68 and into an interior of the airfoil portion of the vane. These gases may degrade the CMC material over time and interrupt heat transfer from the CMC body 62 to the metal body 64 (e.g., platforms 72, 74). To date, known approaches have not fully compensated for the introduction of hot gas into the gap 68. Accordingly, there is a need in the art for improved CMC / metal interfaces that reduce the introduction of hot gases between the materials yet allow for surface and thermal expansion differences therebetween.

SUMMARY

In accordance with one aspect, the deficiencies of the art are addressed by the disclosure of hybrid components as described herein and methods of making the same. The components comprise a first body comprising a metal material and a second body comprising a ceramic matrix composite material. In between the metal and CMC bodies, there is provided a gap. A ceramic fiber material in a suitable form is disposed within the gap. In an aspect and in the case of a gas turbine component, the ceramic fiber material is capable of withstanding the high temperatures associated with gas turbine operation. In addition, the ceramic fiber material is effective to inhibit air flow into the gap between the first body and the second body when a cooling fluid is introduced the component. Still further, the ceramic fiber material transfers a thermal and mechanical load between the first body and the second body, thereby acting as a spring element in certain aspects.

By way of example, upon startup of an associated gas turbine engine which houses such a hybrid component, the rapid temperature increase and high

temperatures will cause the CMC body to expand more quickly than the metal body. The ceramic fiber material (between the CMC body and the metal body) has sufficient deformity to allow the CMC body to grow into it during this stage. Thereafter, the ceramic fiber material stabilizes the interface between the CMC body and the metal body. Upon shutdown of the engine, the CMC body will shrink to a lesser degree than the metal body. In this instance, the ceramic fiber material will expand to accommodate the more rapid shrinkage of the CMC body relative to the metal body.

In another aspect, there is provided a process for forming a component comprising disposing a ceramic fiber material within a gap between a first body comprising a metal material and a second body comprising a ceramic matrix composite material of the component, wherein the ceramic fiber material is effective to inhibit air flow into the gap between the first body and the second body, as well as transfer a thermal and mechanical load between the first body and the second body.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in the following description in view of the drawings that show:

FIG. 1 illustrates a known gas turbine vane having a ceramic matrix composite body and a metal portion (core) extending through the CMC body.

FIG. 2 illustrates a desired but typically infeasible direct interface between the CMC body and the metal portion of the component of FIG. 1.

FIG. 3 illustrates a gap disposed between the CMC body and the metal portion to account for surface defects in either or both of the CMC and metal bodies.

FIG. 4 illustrates an even larger gap disposed between the CMC body and the metal portion to account for thermal expansion differences between the CMC body and the .

FIG. 5 illustrates a cross sectional view of a vane having a gap disposed between the CMC body (shroud portion thereof) and the platforms thereof.

FIG. 6 illustrates a gas turbine engine comprising a hybrid component as described herein in accordance with an aspect of the present invention. FIG. 7 illustrates a hybrid component comprising a ceramic fiber material in between a metallic body and a CMC body in accordance with another aspect of the present invention.

FIG. 8 illustrates another embodiment of a hybrid component comprising a ceramic fiber material in between a metallic body and a CMC body in accordance with another aspect of the present invention.

FIG. 9 illustrates another embodiment of a hybrid component comprising a ceramic fiber material in between a metallic body and a CMC body in accordance with another aspect of the present invention, the metallic body having grooves formed therein to accommodate the ceramic fiber material.

DETAILED DESCRIPTION

Referring again to the figures, FIG. 6 illustrates a gas turbine engine 2 which includes one or more hybrid components formed from a ceramic matrix composite material and a metal material as described herein. The gas turbine engine 2 includes a compressor section 4, a combustor section 6, and a turbine section 8. In the turbine section 8, there are alternating rows of stationary airfoils 18 (commonly referred to as "vanes") and rotating airfoils 16 (commonly referred to as "blades"). Each row of blades 16 is formed by a circular array of airfoils connected to an attachment disc 14 disposed on a rotor 10 having a rotor axis 12. The blades 16 extend radially outward from the rotor 10 and terminate in blades tips. The vanes 18 extend radially inward from an inner surface of vane carriers 22, 24 which are attached to an outer casing 26 of the engine 2. Between the rows of vanes 18 a ring seal 20 is attached to the inner surface of the vane carrier 22. The ring seal 20 is a stationary component that acts as a hot gas path guide between the rows of vanes 18 at the locations of the rotating blades 16. The ring seal 20 is commonly formed by a plurality of ring segments that are attached either directly to the vane carriers 22, 24 or indirectly such as by attachment to metal isolation rings (not shown) attached to the vane carriers 22, 24. During engine operation, high- temperature/high-velocity gases 28 flow primarily axially with respect to the rotor axis 12 through the rows of vanes 18 and blades 16 in the turbine section 8. Referring to FIG. 7 and in accordance with an aspect of the present invention, there is shown a component 100 comprising a first body 102 comprising a metal material 104 and a second body 106 comprising a ceramic matrix composite material 108. One or more gaps 110 are defined between the first body 102 and the second body 106. An intermediate ceramic fiber material 112 (hereinafter ceramic fiber material 112) is disposed within at least one of the gaps 110 between the first body 102 and the second body 106. The ceramic fiber material 112 can withstand the high temperature conditions of the gas turbine engine 2 as a result of utilizing a ceramic material. In addition, the ceramic fiber material 112 is effective to act as a flow deterrent to prevent an air flow from entering the component 100 from an exterior thereof (e.g., hot gases from combustion) that may degrade the metal and/or CMC materials 104, 108 and prevent heat transfer therebetween first body 102 and the second body 106. Still further, the presence of the ceramic fiber material acts 112 as a spring element in that it may transfer a thermal and mechanical load between the first body 102 and the second body 106, and act as a spring depending on the direction of growth/shrinkage of the materials. In particular, the presence of the ceramic fiber material 112 between the first (CMC) body 102 and the second (metal) body 106 will reduce movement of the CMC body 102 (e.g., rattling) during operation or use thereof and thus may reduce the likelihood of structural damage as a result, especially during start up of the associated turbine engine when the component is a gas turbine component.

The component 100 may comprise any suitable component having a metal material 104 and a CMC material 108 as described herein. In certain embodiments, the combination of CMC with metal (e.g., a superalloy) increases a maximum operating temperature of the component by 200-300° C or more relative to the metal component alone. In an embodiment, the component 100 comprises a component for use in the gas turbine engine 2 as shown in FIG. 6, such as one having an operating temperature of 800° C or more, and in certain embodiments of 1200° C or more. In an embodiment, the component 12 comprises a stationary component of a gas turbine, such as a stationary vane or a transition cone. For purposes of illustration, a vane 18 is depicted in FIG. 7. In another embodiment, the component 100 comprises a rotating component for a gas turbine engine, such as a blade 16 (FIG. 6). Alternatively, the component 100 may comprise any other suitable component having a structure as set forth herein.

In addition, it is appreciated that the first body 102 comprising a metal material 104 and the second body 106 comprising a ceramic matrix composite material 108 may each comprise any suitable shape and structure so long as at least one gap 1 10 is defined between the first body 102 and the second body 106. Referring again to FIG. 7, by way of example only, the component 100 comprises a gas turbine vane 18. In this embodiment, the first body 102 comprises an inner platform 1 14 and an outer platform 1 16, and a metal spar 1 18 which extends between the inner platform 1 14 and the outer platform 1 16. The second body 106 comprises an airfoil 120 formed from the CMC material 108 which encompasses the metal spar 1 18 such that the metal spar 1 18 extends through the airfoil 120 to provide a degree of mechanical support for the airfoil 120, as well as allow for transfer of thermal and mechanical loads from the airfoil 120 to the metal spar 1 18. In addition, the second body 106 comprises a shroud portion 122 which transitions the airfoil 120 to each platform surface 124. A gap 1 10 is disposed between the metal spar 1 18 and the airfoil 120 which extends in a radial direction. In addition, there is disposed a gap 1 10 between the inner platform 1 14 and a shroud portion 122 and the outer platform 1 16 and another shroud portion 122. The gaps 1 10 may be of any suitable size for the intended operation of the component 100. In certain embodiments, the size of the gaps 1 10 are sized to allow for manufacturing tolerances and the thermal expansion differences between the components discussed herein, else are kept to a minimum.

In accordance with an aspect, there is provided a ceramic fiber material 1 12 disposed within at least one of the gaps 1 10 between the first body 102 and the body 106. Thus, at least one region of the component 100 comprises a gap 1 10 within which is disposed a ceramic fiber material 1 12. It is appreciated that now all gap locations are necessarily filled the ceramic material 1 12. In an embodiment, as shown in FIG. 7, the vane 1 18 comprises the ceramic material 1 12 between the platforms 1 14, 1 16 and corresponding shroud portions 122, but not in other available gaps 1 10. In the case of a vane 18 in a gas turbine engine 2, the hot gases in contact with the platforms 114,

116 are typically not as hot as the gas that travels along the airfoil 120. Thus, the risk of the ceramic fiber material 112 acting an insulator to too great an extent between the shroud portions 122 and platforms 114, 116 is minimal. In certain embodiments, however, providing the ceramic fiber material 112 in the gaps 110 between the metal spar 118 and airfoil 120, for example, would cause the CMC airfoil 120 to increase to too high a temperature (risk of thermal damage) since the ceramic fiber material 1 12 would act as an insulator. Thus, in certain embodiments, the ceramic fiber material 112 is not disposed between the airfoil 120 and spar 118 but is disposed between the shroud portions 122 and the platforms 114, 116. In further embodiments, the ceramic fiber material 112 may be provided in any other suitable or desired location.

In other embodiments, the ceramic fiber material 112 may be disposed within any and all gaps 110 in the component 100. For example, where hot gas that travels along the airfoil 120 is of a temperature that does not exceed the maximum use temperature of the CMC but is greater than the use temperature of the metal spar 118, the ceramic fiber material 112 may be provided in all gaps 110, including the gaps 110 between the metal spar 118 and airfoil 120. FIG. 8 illustrates a vane 118 as comprising a ceramic fiber material 112 within the gaps 110 between the airfoil 120 and the metal spar 118, as well as between the platforms 114, 116 and corresponding shroud portions 122.

In certain embodiments, the ceramic fiber material 112 is maintained within a respective gap 110 by compression of the first body 102 and/or the second body 106 against the ceramic fiber material 112. In still other embodiments, any other suitable structure may be utilized to assist in retaining the ceramic fiber material 112 within a respective gap 110. For example, as shown in FIG. 9, either or both of the first body 102 and the second body 106 may comprise a groove or channel 132 which is configured to retain at least a portion of the first body 102 and the second body 106 therein.

In accordance with another aspect, to fixedly secure the first body 102 to the second body 104, any suitable structure may be utilized. In an embodiment, an attachment 128 comprises one that limits compression, sliding, or rubbing of the first body 102 vs. the second body 106 is while securing the bodies 102, 106 to one another. In a particular embodiment and as shown in FIG. 7, the attachment 128 comprises one or more clamps 130 which encompasses a portion of the second body 106 and which may secure, fixedly or removably, to the first body 102.

The ceramic fiber material 112 may comprise any suitable ceramic material which will withstand the intended operational temperatures of the component. In an embodiment, the ceramic fiber material 1 12 comprises an oxide ceramic, such as materials available from the Minnesota Mining and Manufacturing Company under the trademark Nextel, including Nextel 720 (alumino-silicate), Nextel 610 (alumina), and Nextel 650 (alumina and zirconia). The fibers may alternatively comprise a non-oxide ceramic material such as silicon carbide available from Dow Corning Corporation under the trademark Sylramic or from the Nippon Carbon Corporation Limited under the trademark Nicalon. The fibers may be in any suitable form. In an embodiment, the ceramic fiber material 112 comprises ceramic fibers (also known as“rovings”) that are oriented into fabrics, filament windings, woven or non-woven mats (e.g., 3D woven mats), braids, or the like. One or more layers of the fibers may be provided within the gaps 110. In a particular embodiment, the ceramic fiber material 112 comprises one or more layers or self-supporting plies 126 of the ceramic fiber material 112 (FIGS. 7-8) .

In a specific embodiment, the ceramic fiber material 112 comprises five or more plies 126 of the fiber material.

The ceramic fiber material 112 is unimpregnated in that the ceramic fiber material 112 is not intentionally pre-impregnated with a ceramic precursor or ceramic matrix material. Typically, a ceramic fiber material utilized in gas turbine components is impregnated with a slurry containing a ceramic precursor or ceramic material prior to (pre-preg) or after lay up of the various fiber layers and then fired at a suitable

temperature. This is done so as to produce a sintered ceramic matrix composite material, which is essentially a fiber reinforced ceramic matrix. The ceramic precursor or ceramic matrix material may comprise a similar type of material as the reinforcing fibers, such as an oxide or non-oxide material. In the present invention, however, the ceramic fiber material 112 is not impregnated with a ceramic or ceramic precursor material. In this way, the ceramic fiber material 112 is able to compress and relax in response to mechanical and thermal stresses and transfer loads between the first body 102 and the second body 106.

With respect to the metal material 104, the metal material 104 of the first body 102 may comprise any suitable metal (e.g., alloy) material suitable for its intended use in the component 100, e.g., to provide a vehicle for mechanical and/or thermal load transfer from the second body 106 of CMC material 108. In an embodiment, the metal material 24 comprises a superalloy material, such as a nickel-based or a cobalt-based superalloy material, as is well known in the art. The term "superalloy" may be understood to refer to a highly corrosion-resistant and oxidation-resistant alloy that exhibits excellent mechanical strength and resistance to creep - even at high

temperatures. Exemplary superalloy materials are commercially available and are sold under the trademarks and brand names Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939), Rene alloys (e.g. Rene N5, Rene 41 , Rene 80, Rene 108, Rene 142, Rene 220), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718, X-750, ECY 768, 262, X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystal alloys, GTD 111 , GTD 222, MGA 1400, MGA 2400, PSM 116, CMSX-8, CMSX-10, PWA 1484, IN 713C, Mar-M- 200, PWA 1480, IN 100, IN 700, Udimet 600, Udimet 500 and titanium aluminide, for example. The first body 102 of metal material 104 may be formed by any suitable process known in the art, such as by casting or an additive manufacturing process.

With respect to the CMC material 108, the CMC material 108 of the second body 106 may comprise any suitable ceramic or a ceramic matrix material which hosts a plurality of reinforcing fibers as is known in the art. For example, the CMC material 108 may comprise a fiber reinforced matrix material or metal reinforced matrix material as may be known or later developed in the art, such as one commercially available from the COI Ceramics Co. under the name AS-N720. If a fiber reinforced material is used, similar to the ceramic fiber material 112, the fibers may comprise oxide ceramics, non- oxide ceramics, or a combination thereof. For example, the oxide ceramic fiber composition can include fibers commercially available from the Minnesota Mining and Manufacturing Company under the trademark Nextel, including Nextel 720 (alumino- silicate), Nextel 610 (alumina), and Nextel 650 (alumina and zirconia). For another example, the non-oxide ceramic fiber composition can include those commercially available from the COI Ceramics Company under the trademark Sylramic (silicon carbide), and from the Nippon Carbon Corporation, Limited under the trademark

Nicalon (silicon carbide).

The matrix material composition that surrounds the fibers may be made of an oxide or non-oxide material, such as alumina, mull ite, aluminosilicate, ytrria alumina garnet, silicon carbide, silicon nitride, silicon carbonitride, and the like. The CMC material 108 may combine a matrix material with fibers of a different composition (such as mullite/silica) or the matrix material and fibers may be of the same composition (alumina/alumina or silicon carbide/silicon carbide). The fibers may be continuous or long discontinuous fibers, and may be oriented in a direction generally parallel, perpendicular, or otherwise disposed relative to the major length of the CMC material 108. The matrix composition may further contain whiskers, platelets, particulates, fugitives, or the like. The reinforcing fibers may be disposed in the matrix material in layers, with the plies of adjacent layers being directionally oriented to achieve a desired mechanical strength.

The fibers may be provided in various forms, such as a woven fabric, blankets, unidirectional tapes, rovings, plies, and mats, e.g., 3D woven mats. In an embodiment, the CMC fibers may be any material which may be utilized for the ceramic fiber material 112 (and vice-versa), but the fibers of the CMC material 108 are also impregnated with a ceramic or ceramic precursor material and subjected to a suitable heat treatment process, e.g., sintering. A variety of techniques are known in the art for making a CMC material and such techniques can be used in forming the CMC material 108 for use herein. In addition, exemplary CMC materials 108 are described in U.S. Patent Nos. 8,058,191 ; 7,745,022; 7,153,096; 7,093,359; and 6,733,907, the entirety of each of which is hereby incorporated by reference. In a particular embodiment, the CMC material 108 may be manufactured by a process as disclosed in PCT/US2016/059029, the entirety of which is incorporated by reference herein, wherein a ceramic material is injected into a fiber material to form a ceramic fiber composite which is then 3D printed in a desired pattern and heat treated to form the desired CMC material 108.

In accordance with another aspect, there is provided a method for making a component 100 as described herein. First, a first body 102 comprising a metal material 104 and a second body 106 comprising a CMC material 108 are provided. The first body 102 may be formed by any suitable process, such as a casting process, an additive manufacturing process, or any other process known in the art. When an additive manufacturing process is utilized for the metal first body 102, the additive manufacturing process may comprise metal injection molding, selective laser

manufacturing or melting (SLM), a selective laser sintering (SLS), a laser metal deposition technique, or any other additive manufacturing process. SLM and SLS are manufacturing techniques that build components layer by layer from powder beds. In these processes, a powder bed of a component final material, or a precursor material, is deposited onto a working surface. Thereafter, laser energy is directed onto the powder bed following a cross-sectional area shape of the component to create a layer or slice of the component. The deposited layer or slice then becomes a new working surface for the next layer. The second body 106 of the CMC material 108 may also be formed by any suitable process, such as those including a typical layup process, the stacking of CMC laminates, or via 3D printing a ceramic loaded fiber material into the desired form. Alternatively, the first or second bodies 102, 106 may be provided from a commercially available source.

Once provided, the method comprises providing the ceramic fiber material 112 within a gap 110 between the first body 102 and the second body 106. This may be done by any suitable method, such as by laying up the ceramic fiber material 112, e.g., plies 112, against the first body 102, for example, and then positioning the second body 106 against the ceramic fiber material 112. The second body 106 may then be secured to the first body 102 by any suitable process, such as by positioning the second body 106 within the clamp 130 and tightening down the clamp 130 on the second body 106.

In certain embodiments, by securing the second body 106 and the first body 102 to one another (or by any other suitable tensioning method), the ceramic fiber material 112 is compressed. The inventors have surprisingly found that compressing the fiber material prior to subjecting the component 100 to high temperature operation of the gas turbine 112 may further aid the ability of the fiber material 112 to expand or contract as needed to allow for load transfer and differential thermal growth between the metal material 104 and the CMC material 108. Once assembled, the component may be utilized in high temperature operation.

While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.

Claims

CLAIMS What we claim is:

1. A component (100) comprising:

a first body (102) comprising a metal material (104);

a second body (106) comprising a ceramic matrix composite material (108); a gap (110) disposed between the first body (102) and the second body (106); and

a ceramic fiber material (112) disposed within the gap (110), the ceramic fiber material (112) effective to inhibit air flow into the gap (110) between the first body (102) and the second body (106), and transfer a thermal and mechanical load between the first body (102) and the second body (106).

2. The component (100) of claim 1 , wherein the ceramic fiber material (112) comprises a plurality of plies (126) of ceramic fibers.

3. The component (100) of claim 1 , wherein the ceramic fiber material (112) comprises an oxide-oxide material.

4. The component (100) of claim 1 , wherein the component (100) comprises a hot gas path component (115) of a gas turbine engine (2).

5. The component (12) of claim 4, wherein the component (100) comprises a gas turbine vane (115).

6. The component (100) of claims 1 to 5,

wherein the first body (102) comprises a metal platform (114, 116);

wherein the second body (106) comprises a ceramic matrix composite shroud portion (122), the shroud portion (122) secured to the metal platform (114, 116); and wherein the ceramic fiber material (112) is disposed between the shroud portion (122) and the metal platform (104, 116).

7. The component (100) of claim 6, wherein the second body (106) further comprises a ceramic matrix composite airfoil (120) and the first body (102) comprises an inner platform (1 14), an outer platform (1 16), and a spar (1 18) extending radially through the airfoil (120) and between the inner platform (1 14) and the outer platform (1 16), and wherein the ceramic fiber material (1 12) is disposed between the shroud portion (122) and the metal platform (104, 1 16).

8. The component (100) of claim 1 , wherein the first body (102) or the second body (106) comprises a channel (132) formed therein and dimensioned to accommodate the ceramic fiber material (1 12) therein.

9. A process for forming a component (100) comprising:

disposing a ceramic fiber material (1 12) within a gap (1 10) between a first body (102) comprising a metal material (104) and a second body (106) comprising a ceramic matrix composite material (108) of the component (100), wherein the ceramic fiber material (1 12) is effective to inhibit air flow into the gap (1 10) between the first body (102) and the second body (106), and transfer a thermal and mechanical load between the first body (102) and the second body (106).

10. The process of claim 9, further comprising compressing the ceramic fiber material (1 12) within the gap (1 10) prior to subjecting the ceramic fiber material (1 12) to operating conditions of a gas turbine engine (2).

1 1 . The process of claim 9, wherein the ceramic fiber material (1 12) comprises an oxide-oxide material.

12. The process of claim 9, wherein the component (100) comprises a gas turbine vane (10).

13. The process of claim 12,

wherein the first body (102) comprises a metal platform (114, 116);

wherein the second body (106) comprises a ceramic matrix composite shroud portion (122), the shroud portion (122) fixedly secured to the metal platform (114, 116); and

wherein the ceramic fiber material (112) is disposed between the shroud portion (122) and the metal platform (104, 116).