WO2019209267A1

WO2019209267A1 - Ceramic matrix composite component and corresponding process for manufacturing

Info

Publication number: WO2019209267A1
Application number: PCT/US2018/029117
Authority: WO
Inventors: Zachary D. Dyer; David Gamblin
Original assignee: Siemens AG; Siemens Corp
Current assignee: Siemens AG; Siemens Corp
Priority date: 2018-04-24
Filing date: 2018-04-24
Publication date: 2019-10-31
Anticipated expiration: 2020-10-24

Abstract

There is provided a component (110) comprising a plurality of ceramic matrix composite members (111) stacked on one another to define an airfoil (116). The airfoil (116) has a radial axis (130) extending therethrough and a horizontal axis (132) perpendicular to the radial axis (132). The ceramic matrix composite members (111) each having a degree of curvature (119) and stacked on one another non-parallel to the horizontal axis (132).

Description

CERAMIC MATRIX COMPOSITE COMPONENT AND CORRESPONDING

PROCESS FOR MANUFACTURING

FIELD

The present disclosures relates to high temperature components, and more particularly ceramic matrix composite (CMC) components formed from a plurality of ceramic matrix composite members stacked in an off-horizontal axis orientation and having a curved profile, as well as methods for making the same.

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 high temperature and high pressure gas. This working gas then travels past the combustor transition piece and into the turbine section of the turbine.

Generally, 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. High efficiency of a 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 turbine components from extreme temperatures such as 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. State of the art superalloys with additional protective coatings are commonly used for hot gas path components of gas turbines. In view of the substantial and longstanding development in the area of superalloys, however, it figures to be extremely difficult to further increase the temperature capability of superalloys.

Accordingly, ceramic matrix composite (CMC) materials have been developed with greater temperature resistance relative to superalloys. Such CMC materials may include a ceramic or a ceramic matrix material, either of which hosts a plurality of reinforcing fibers. It has been found, however, that forming turbine components from CMC materials may be challenging due to, amongst other things, the difficulty in orientating fibers at edges of the component in the complex shapes typical of many turbine components. In addition, though the fibers provide a degree of added mechanical strength to the structure, CMC materials are known to have weak interlaminar strength, and are thus prone to delamination or failure upon exposure to significant mechanical or thermal loads.

CMC structures may be formed by a variety of techniques. For one, CMC structures may be formed by stacking (laying up) a plurality of fibrous members (e.g., plies) on one another in a horizontal stack (akin to a stack of paper on a table). In some embodiments, the fibrous members are pre- impregnated with a ceramic or ceramic precursor material. In other embodiments, the fibers are laid down and then impregnated with the material. The plies are then sintered to provide the desired component. In other embodiments, to form a CMC structure, CMC laminates are positioned on one other in a horizontal stack to form the desired component, wherein each laminate plate is made from a plurality of such plies or fibrous members. The stacked laminates allow for increased control of fiber orientation and allow for more complex shapes to be formed. In any case, the formed CMC component often requires additional cooling for use at the high temperatures that would produce desirable engine efficiency, particularly due to the CMC’s relatively low thermal conductivity. Unfortunately, directing a coolant flow through the CMC component often results in a loss of cooling fluid through the weak interlaminar portions of the horizontally stacked fibrous structures, particularly in the case of CMC laminate structures. The loss of cooling fluid and pressure reduces cooling efficiency, and thus the operating temperatures that the component may be safely operated in without causing the aforementioned delamination or other structural damage (such as embrittlement, cracking, or the like, of the component). Accordingly, there is a need for CMC structures and processes for making the same having increased interlaminar strength and reduced cooling air loss there through.

SUMMARY

Aspects of the present invention are directed to components and processes for making ceramic matrix composite (CMC) components formed from a plurality of CMC members such as plies, laminates, or the like, which are stacked (laid up)“non-horizontally” (e.g., vertically or parallel to a radial axis of the component) as opposed to the typical horizontal orientation of such CMC members in a component. The non-horizontal orientation of the CMC members substantially prevents or reduces interlaminar separation of the CMC members. This, in turn, extends component lifetime, reduces cooling air loss. In addition, the components described herein may have vastly different mechanical and thermal resistance profiles relative to conventionally formed CMC components - due to their vast differences in main fiber axis alignment.

In certain embodiments, the reduced need for cooling air, improved

mechanical strength, and improved thermal resistance relative to

conventionally formed components allow for higher temperature operation. In the case of a gas turbine component, such higher temperatures and efficient cooling may lead to increased efficiency within the associated gas turbine. In accordance with an aspect, there is provided a component comprising a plurality of ceramic matrix composite members arranged (e.g. , stacked) on one another to define an airfoil. The airfoil has a radial axis extending therethrough and a horizontal axis perpendicular to the radial axis. The ceramic matrix composite members each have a degree of curvature and are stacked on one another non-parallel to the horizontal axis.

In accordance with another aspect, there is provided a process for manufacturing a gas turbine component. The process comprises stacking a plurality of ceramic matrix composite members non-horizontally on one another to define a precursor block comprising the ceramic matrix composite members; and heating the precursor block for a duration and at a temperature effective to sinter the precursor block. In an embodiment, the precursor block defines an airfoil. In other embodiments, the precursor block defines a body having a larger volume or size than the airfoil to be manufactured. In this case, the process may further comprise shaping the precursor block to form an airfoil, wherein the airfoil has a radial axis extending therethrough and a horizontal axis perpendicular to the radial axis, the ceramic matrix composite members each having a degree of curvature and stacked on one another non- parallel to the horizontal axis.

BRIEF DESCRIPTION OF THE DRAWINGS

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

FIG. 1 illustrates a gas turbine in accordance with an aspect of the present invention.

FIG. 2 illustrates a blade formed by a process in accordance with an aspect of the present invention.

FIG. 3 illustrates a cross-section of an embodiment of an airfoil formed from a plurality of non-horizontally oriented plies taken at line A-A of FIG. 2 in accordance with an aspect of the present invention. FIG. 4 illustrates a cross-section of an embodiment of an airfoil formed from a plurality of non-horizontally oriented laminates taken at line A-A of FIG. 2 in accordance with an aspect of the present invention.

FIG. 5 illustrates a mean camber line of an airfoil formed by a process in accordance with an aspect of the present invention.

FIG. 6 illustrates an embodiment of another airfoil formed in

accordance with an aspect of the present invention.

FIG. 7 illustrates an airfoil having a plurality of cooling channels formed therein in accordance with an aspect of the present invention.

FIG. 8 illustrates the formation of a cooling channel via two adjacently stacked laminates in accordance with an aspect of the present invention.

FIG. 9 illustrates a step in the manufacture of a component as described herein, wherein a plurality of plies are assembled together to form a precursor block from which a desired airfoil is shaped in accordance with an aspect of the present invention.

FIG. 10 illustrates a step in the manufacture of a component as described herein, wherein a plurality of laminates are assembled together to form a precursor block from which a desired airfoil is shaped in accordance with an aspect of the present invention.

FIG. 11 illustrates a stationary vane form by a process describe herein in accordance with an aspect of the present invention.

DETAILED DESCRIPTION

Now referring to the figures, FIG. 1 illustrates a known gas turbine engine 2 having 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 (not shown) 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.

In accordance with an aspect, one or more of the components of the gas turbine 2 are formed from a plurality of curved and non-horizontally oriented ceramic matrix composite (CMC) members. As shown in FIG. 2, for example, there is shown an exemplary component 1 10, e.g., airfoil 1 16, comprising a plurality of the CMC members 1 1 1 . As shown, the airfoil 1 16 comprises a body 1 17 having a leading edge 1 18 and a trailing edge 120 which extends between a root 122 and a tip 124 of the airfoil 1 16. In addition, opposed sidewalls of a pressure side 126 and a suction side 128 are defined between the leading edge 1 18 and the trailing edge 120. The airfoil 1 16 thus includes a leading edge 1 18, a trailing edge 120, opposed pressure and suction sidewalls 126, 128 extending along a radial axis 130 from the root 122 towards the tip 124. As shown, a horizontal axis 132 extends perpendicular to the radial axis 130 at any point along the radial axis 130. In addition, the horizontal axis 132 may extend in any direction about a circumference (C) of the radial axis as shown. In certain embodiments, the radial axis 130 represents a longest dimension of the airfoil 1 16 and a longest dimension of the CMC members 1 1 1 are oriented“non-horizontally.”

In accordance with an aspect of the present invention, the CMC members 1 1 1 are stacked on one another“non-horizontally” as opposed to the horizontally in conventional CMC components. Conventional CMC laminates or fibrous structures (e.g., plies) stacked horizontally on one another from the root 122 towards the tip 124 - parallel to the horizontal axis 132. In contrast to the conventional structures, the CMC members 111 descried herein are not stacked one after the other from the root 122 toward the tip 124, but instead are stacked“non-horizontally” such that the CMC members 111 are each stacked on one another non-parallel to the horizontal axis 132. Put another way, by“non-horizontally,” it may be said that the CMC members 111 are stacked in a direction from the pressure side 126 to the suction side 128 (or vice-versa from suction side 128 to pressure side 126) to build the component 110 versus stacking from the root 122 towards the tip 124 of the component 110 as in conventional structures. In particular embodiments, the CMC members 111 are stacked vertically. By“vertically, it is meant that the CMC members 111 are each stacked on one another at an angle of 90 degrees relative to the horizontal axis 132, or parallel to the radial axis 130.

The present inventors have found that this unique, non-conventional orientation of the CMC members 111 provides for distinct fiber orientations not yet available to date which may have substantial additional thermal and mechanical strength benefits. For one, the nonconventional (e.g., vertical) orientation of the CMC members 111 place the members 111 in a position where cooling air leakage between adjacent laminates is substantially less likely to occur (when cooling air is flowed through the structure). In addition, the nonconventional orientation further provides the component 110 with a substantially increased interlaminar strength, typically a weakness of known CMC materials, due to the fact there are no structures (e.g., laminates) oriented horizontally that are prone to delamination / separation from one another.

The CMC members 111 as described herein may comprise any suitable fiber material, whether woven or non-woven and in any suitable form, which is impregnated with or otherwise incorporated within a ceramic or a ceramic precursor material (hereinafter“ceramic material”). In an embodiment, the CMC members 111 comprise ceramic fibers (also known as “rovings”) that are oriented into fabrics, filament windings, braids, or the like. Referring to FIG. 3, in an embodiment, the CMC members 111 may comprise a plurality of plies 115 which are stacked on one another in a non-horizontal direction. FIG. 3 is an embodiment of a cross-section of the airfoil 116 of FIG. 2 taken at line A-A. Each ply 115 comprises a layer of fibers in a desired form which are individually or collectively impregnated with a ceramic or ceramic precursor material. In certain embodiments, the plies are self-supporting. In certain embodiments, the plies 115 are provided in the form of“pre-preg” having a fiber material already impregnated with a ceramic or ceramic precursor material, which are then oriented vertically as described herein.

In other embodiments and as shown in FIG. 4, the CMC members 111 may instead comprise a plurality of CMC laminates 112. Individual CMC laminates 112 are typically formed from a plurality of plies that are

impregnated with a ceramic or ceramic precursor material and sintered to form the individual laminates 112. It is appreciated that CMC members 111 (whether laminates, plies, or other form) utilized to form the component 110 may comprise any suitable size, shape, and number necessary to form the desired component.

In addition to the non-horizontal orientation described herein, the CMC members 111 also comprise a degree of curvature. The degree of curvature not only provides for components 100 with unique strength and temperature profiles not yet achieved, but also simplifies manufacturing of a component formed from the CMC members 111. Referring again to FIG. 4, there is shown a cross-section of the airfoil 116 of FIG. 2 taken at line A-A. As shown, in this embodiment, an airfoil 116 may be formed from a plurality of laminates 112 having a degree of curvature or a curved shape 119 such that when stacked non-horizontally (e.g., vertically), the members 111 collectively define the airfoil 116. In certain embodiments, each member 111 comprises a degree of curvature 119 that follows a mean camber line 134 of the airfoil 116 to be formed by the members 112. For purposes of illustration, FIG. 5 shows an airfoil 116 having a mean camber line 134 that extends between the leading edge 118 and the trailing edge 120, and includes a locus of points equidistant between an upper surface 136 and a lower surface 138 of the airfoil 116.

The present inventors have found that the components as described herein may provide differing mechanical strength, thermal conductivity, or other differing physical characteristics depending on the orientation of the members 111 relative to the horizontal axis 132. In certain embodiments, this may be due to the different fiber orientation of the ceramic matrix composite (CMC) material in the laminates previously unavailable without this off- horizontal axis alignment. Thus, in accordance with an aspect of the present invention, the angle of the CMC members 111 relative to the horizontal axis 132 may be varied to achieve a particular mechanical strength or high temperature resistance profile. As mentioned, it is also contemplated that the members may further be oriented at any desired position about a

circumference (C) about the radial axis 130 as shown in FIG. 2.

As mentioned, the CMC members 111 may be stacked“non- horizontally,” or at any angle other than parallel or 0° relative to the horizontal axis 130 to build the component 110. In an embodiment, the CMC members 111 are stacked at an angle from 10 degrees to about 90 degrees relative to the horizontal axis 132. In specific embodiments, the CMC members 111 are stacked vertically or parallel to or along the radial axis 130. Put yet another way, in certain embodiments, the CMC members 111 may be stacked on one another at an angle of 90 degrees relative a horizontal axis 132 extending through the airfoil 115.

To reiterate, FIG. 4 illustrated an embodiment wherein the fibrous members 111 comprise a plurality of laminates 112. In other embodiments, the members 111 may comprise any other suitable structure having a fiber reinforced ceramic matrix. For example, as was shown in FIG. 3, the members 111 each comprised a plurality of plies 115 of the fiber material, which are impregnated individually (prior to lay up) or collectively with a ceramic material and subjected to a suitable heat treatment, e.g., sintering, process to form the component 110.

Regardless of the form (plies, laminates, or otherwise) of the CMC members 111 , upon sintering, the CMC members 111 each comprise a CMC material. The CMC material may comprise any suitable fiber reinforced matrix material, such as one commercially available from the COI Ceramics Co. under the name AS-N720. If a fiber reinforced material is used, the fibers may comprise oxide ceramics, non-oxide ceramics, or a combination thereof. For example, the oxide ceramic fiber composition can include those

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, mullite, aluminosilicate, ytrria alumina garnet, silicon carbide, silicon nitride, silicon carbonitride, and the like. In an embodiment, the CMC material comprises an oxide-oxide material (oxide fibers and oxide matrix). The CMC material may combine a matrix composition with a reinforcing phase of a different composition (such as mullite/silica), or may be of the same composition (alumina/alumina or silicon carbide/silicon carbide). The fibers may be continuous or long discontinuous fibers. The matrix composition may further contain whiskers, platelets, particulates, or fugitives, or the like. In addition, the reinforcing fibers may be disposed in the matrix material in layers, with the plies of adjacent layers being individually and directionally oriented to achieve a desired mechanical strength.

The fibers may be provided in various forms, such as a woven fabric, blankets, unidirectional tapes, and mats. As was mentioned, the fibers are infused with a ceramic material (prior to or after lay up) and subjected to a suitable heat treatment, e.g., sintering, process to provide the CMC members 111. techniques are known in the art for making a CMC material and such techniques can be used in forming the CMC material of the members 111. In addition, further exemplary CMC materials and processes for making the same 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 still another embodiment, the individual CMC members 111 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 to form individual CMC members 1 1 1 .

The component 110 may comprise any desired component formed from the plurality of CMC members 111 in the orientation(s) described herein. In an embodiment, the component 110 may comprise a turbine component as is known in the art for use at high temperatures. In a particular embodiment, the component 110 may comprise a gas turbine component configured for use in a combustor turbine hot gas section. By way of example, the component 110 may comprise a stationary part of a gas turbine engine, such as a transition duct, an exhaust cone, a vane 18 (FIG. 1 ), or the like. In other embodiments, the component may comprise a rotating part of a gas turbine, such as a blade 16 (FIG. 1 ). Thus, in certain embodiments, the airfoil 116 described herein may be mounted on a platform as is known in the art at the root and/or the tip thereof.

In accordance with another aspect, 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, in certain embodiments and as shown in FIG. 6, the airfoil 116 may further comprise at least one metal support 140 which extends radially through a body thereof. In certain embodiments, more than one metal support 140, e.g., two or more, may be provided extending radially through the airfoil 116. In some embodiments, the metal support 140 may be pre-formed. In other embodiments, the metal support 140 may be formed by an additive manufacturing process. In this way, the metal support 140 may comprise an optimal interface between the CMC material of the members 111 and the metal support 140 along an entire radial length of the airfoil 116. Exemplary additive manufacturing processes for producing a support through an airfoil 116 are described in PCT Application No.

PCT/US2015/023017, the entirety of which is hereby incorporated by reference.

The metal support 140 may comprise any suitable metal material which will provide an added strength to the members 111 , as well as allow for cooling of the CMC material thereof by being in contact therewith or by being in close proximity thereto such that the CMC material transfers heat to the metal support 140. In certain embodiments, the metal material may comprise a superalloy material, such as a Ni-based or a Co-based superalloy material as are 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.

In accordance with another aspect, the CMC members 11 may further include any suitable structure that facilitates positioning of a first CMC member 111 on a second CMC member 111 as the members 111 are stacked on one another. For example, in an embodiment, the CMC members 1 1 1 comprise CMC laminates 1 12, and a first laminate may comprise an indent which is dimensioned to receive a tab from a corresponding second laminate. Alternatively, any other suitable structure may be utilized for maintaining or facilitating the positioning, interlocking, and/or stacking of the members 1 1 1.

In accordance with another aspect, the component 1 10 may further comprise one or more cooling passages that extend in any desired direction and pattern through the component. For example, FIG. 7 illustrates component 1 10 as further comprising a plurality of cooling passages 142 formed within an interior of the body 1 17 of airfoil 1 16 such that a cooling fluid may be flowed through the cooling passages 142 to draw heat from the body 1 17 of the component 1 10 by convection during high temperature operation. The cooling passages 142 may have any suitable configuration and spacing throughout the CMC body 1 17 such that the cooling passages 142 provide a desired cooling effect through the airfoil.

In an embodiment, at least a portion of the cooling passages 142 extend in a lateral direction from the leading edge 1 18 toward the trailing edge 120 (or vice-versa) of the CMC body 1 17. In addition, the cooling passages 142 may include those that extend in a direction from the root 122 to the tip 124 or along the radial axis 130 of the CMC body 1 17. The radial and lateral extending cooling passages 142 may intersect as needed to provide for a flow of a cooling fluid, e.g., air, through the stacked laminate structure. In certain embodiments, the cooling passages 142 comprise a serpentine arrangement throughout the CMC body 1 17. Further, in certain embodiments, at least some of the cooling passages 142 have an outlet at the trailing edge 120 of the CMC body 1 17 to allow for exit of the cooling air (with extracted heat).

In certain embodiments, the cooling passages 142 may be formed in the airfoil 1 16 by drilling or otherwise machining the passages 142 within the CMC body 1 17 following sintering thereof. In certain embodiments, such as when laminates 1 12 are utilized, the cooling passages 142 may be defined by one or more channels formed into a depth of the laminates 112 such that when the members 111 are stacked on one another, the cooling passages 142 are defined therein. For example, FIG. 8 illustrates channels 144 in abutting laminates 112 overlap one another to collectively define a cooling passage 142. In certain embodiments, the cooling passages 142 may further include fins, stanchions, buttons, or the like associated therewith to increase the heat transfer properties of the channels.

To allow for delivery of a cooling fluid, e.g., air, to the cooling passages 142, at least some of the channels define an inlet in fluid communication with a source of a cooling fluid such that the cooling fluid may be flowed

throughout the body 117. In an embodiment, for example, the inlet(s) are in fluid communication with an air compressor to deliver an amount of air through the passages 142. In certain embodiments, the inlet comprises an opening in the body 117 and in fluid connection with the passages 142 such that a cooling fluid flowed into the opening will travel into respective cooling passages 142. It is appreciated that the shape, number, and orientation of the plies may be varied as desired to achieve different configurations, as well as different mechanical and thermal characteristics.

In accordance with another aspect, there is provided a process for manufacturing“off-horizontal axis” CMC components 110 as described herein. In certain embodiments, the CMC members 111 may have a size and shape, and may be stacked non-horizontally as described herein that the CMC members 111 define a“precursor block 150” that is larger than the desired dimensions of the component 110. Upon sintering, the desired airfoil 116 may be shaped from the larger precursor block 150 by any suitable physical, chemical, or other method, such as by grinding down the precursor block 150 to a desired shape and desired dimensions. In other embodiments, the CMC members 111 collectively define a desired shape of the airfoil 116 without shaping. In this case, the precursor block 150 defines the airfoil 116 as described herein after sintering. Referring again to FIG. 5, there was shown a cross-section of an airfoil 116 having a mean camber line 134 extending from the leading edge 118 to the trailing edge 120, which represents the final desired product. In an embodiment, the airfoil 116 represents a desired finished product. To manufacture the airfoil 116, FIG. 9 illustrates the laying down of CMC members 111 , e.g., plies 115, in an lower shell of a mould 146 having an inner surface 148 configured for housing a plurality of the CMC members 111 therein. In an embodiment, the inner surface 148 has a contour that substantially follows a desired contour of an outer surface of the desired airfoil 116. In a particular embodiment, embodiment, the inner surface 148 has a contour which follows the mean camber line 134 of the desired airfoil 116 to be formed.

Typically, the mould 146 includes a heat source for sintering the plurality of CMC members 111 and forming the precursor block as will be discussed below. In addition, in certain embodiments, the mould 146 may include a source of a ceramic or ceramic precursor material to impregnate the plurality of CMC members 111 with a ceramic or ceramic matrix material (if needed). Further, the mould 146 may include a vacuum source or any other suitable structure for maintaining the plies 115 in a fixed position as they are laid down in the mould 146.

FIG. 9 illustrates the CMC members 111 laid down in the mould 146 one after the other to define a precursor block 150 having dimensions larger than the dimensions of the desired airfoil 116. In the embodiment of FIG. 9, the CMC members 111 are illustrated as being in the form of plies 115 as described herein, however, understood that the present invention is not so limited. In other embodiments, the CMC members 111 may comprise laminates 112 as described herein or any other structure which comprises a fiber material impregnated with an amount of a ceramic or ceramic precursor material or fiber-reinforced ceramic material. FIG. 10 illustrates a plurality of laminates 112 stacked on one another form the precursor block 150. In any case, the CMC members 111 are stacked on one another such that the members 111 extend in an off-horizontal direction (e.g., vertical direction or out of the page). In addition, the CMC members 111 may be maintained in a fixed position within the mould 148 as they are laid down by any suitable method or structure, such as via vacuum, compression, or the like.

Once the precursor block 150 is formed and are in a fixed position in the mould 146, the precursor block 150 may be subjected to a suitable heat treatment or sintering process. A fully sintered precursor block 150 is illustrated in FIG. 10. It is appreciated that upon sintering, the visible division between adjacent plies, laminates, or the like in the precursor block 150 may be less visible to the naked eye. The sintering process comprises subjecting the CMC members 111 in the desired form for a duration and at a

temperature sufficient to rigidize the precursor block of CMC members 111 and stabilize the ceramic matrix composite (CMC) material therein against liquid water. In one embodiment, the heating comprises heating the CMC members 111 at a temperature(s) greater than about 500° C. In an

embodiment, the heating is done for a duration of from 10 minutes to 24 hours, for example, and may be done isothermally or with a temperature gradient.

In certain embodiments, in a next step following sintering, the precursor block 150 may be shaped into the final desired component shape (see e.g., airfoil 116 in dotted lines in FIG. 10). In accordance with an aspect, prior to shaping, the precursor block 150 may be rotated in one or more dimension (relative to an x, y, z axis thereof) in order to provide further“off axis” orientation of the CMC members 111 within the airfoil 116. The shaping may be done by any suitable process, such as mechanical process, such as one that removes material and shapes the precursor block 150 into its final form. The shaping may be done by any suitable process in the art such as by a suitable machining or water jetting process. In this way, the shaping removes any material necessary to leave behind a component 110, e.g., an airfoil 116, of a desired dimension. Following sintering and before or after shaping, any further processing steps may be performed to form the desired component 1 10. For example, in certain embodiments, the component 1 10 may be provided with an

overwrapping comprising one or more plies of a CMC material and having a desired thickness. In the case of laminates 1 12, the overwrapping may limit excessive movement between the laminates 1 12 in the component 1 10. In addition and as previously discussed, cooling channels may be formed within the body 1 17 of the airfoil 1 16 by machining, mating channels within the CMC members 1 1 1 , or the like. An airfoil with cooling channels formed therein was illustrated in FIG. 7.

In still other embodiments, it may be desirable to further apply an outer thermal barrier coating (TBC) to the body 1 17 of the airfoil 1 16 once formed.

A TBC 152 is shown generally in FIG. 6 on an outer surface of the airfoil 1 16. In an embodiment, the deposition of the TBC 152 is done following sintering. TBCs are well known in the art and may be applied to the outer surface of the airfoil 1 16 by any suitable process, e.g., a thermal spray process, a slurry- based coating deposition process, or a vapor deposition process. The TBC 152 may comprise any suitable material which provides an increased temperature resistance to the body when applied to a surface thereof.

In an embodiment, the TBC 152 comprises a stabilized zirconia material. For example, the TBC 152 may comprise an yttria-stabilized zirconia (YSZ), which includes zirconium oxide (ZrC ) with a predetermined concentration of yttrium oxide (Y2O3). In another embodiment, the TBC 152 may comprise a magnesia stabilized zirconia, ceria stabilized zirconia, aluminum silicate, or the like.

In yet another embodiment, the TBC 152 may comprise a pyrochlore structure. In an embodiment, the pyrochlore structure has the empirical formula A2B2O7 or in general terms A_vB_xO_z where v = 2, x = 2 and z = 7. Deviations from this stoichiometric composition for v, x and z may occur as a result of vacancies or minor, deliberate or undeliberate doping. In the formula A_vBxOz where v = 2, x = 2 and z = 7, gadolinium (Gd) is typically used for A, and hafnium and/or zirconium (Hf, Zr) are typically used for B. In this case too, minor deviations from this stoichiometry may occur. When Hf and Zr are used as B (e.g., Gdv(Hf_xZr_y)Oz), x + y = 2. In other embodiments, B = Zr or Hf individually, and the pyrochlore structure may comprise one of gadolinium zirconate (Gd₂Zr₂07) or gadolinium hafnate (Gd₂Hf₂07). In still other embodiments, the TBC 36 may comprise a bilayer 8YSZ/59 weight percent gadolinium stabilized zirconia (8YSZ/59GZO) coating, a bilayer 8YSZ/30-50 weight percent yttria stabilized zirconia (“30-50 YSZ”) coating, or the like.

In yet another embodiment, the TBC 152 may comprise a

dimensionally stable, abradable, ceramic insulating material comprising a plurality of hollow ceramic particles dispersed therein. The hollow particles may be of any suitable dimension, and in one embodiment may be from 1 -100 micron in diameter. The TBC 152 may be applied by any suitable process, such as a thermal spray process, a slurry-based coating deposition process, or a vapor deposition process as is known in the art. In addition, the TBC 152 may further comprise a degree of porosity suitable for the desired application. Further, the TBC 152 may be of any suitable thickness for its intended use, such as from 0.1 to 2.0 mm.

In certain embodiments, a bond coat (not shown) is known in the art is further provided in order to improve adhesion of the TBC 152 to its underlying surface (body 12). The material for the bond coat may comprise any suitable material. For example, an exemplary bond coat layer may comprise an MCrAIY material, where M denotes nickel, cobalt, iron, or mixtures thereof, Cr denotes chromium, Al denotes aluminum, and Y denotes yttrium. In other embodiments, the bond coat may comprise alumina, yttrium aluminum garnet (YAG), or other suitable ceramic-based material, e.g., a rare earth oxide and/or rare earth garnet material. The bond coat may also be applied by any known process, such as via a thermal spraying or a slurry-based deposition process. When the component 1 10 comprises an airfoil 1 16 as shown in FIG. 1 1 , once the airfoil 1 16 has been formed, the airfoil 1 16 may be properly associated with one or more platforms 154 by any suitable process to complete manufacture of the component 1 10.

The above-described components and systems describe novel and inventive components and processes for making the same that introduce numerous new design possibilities for CMC components. For one, the“off- horizontal axis” CMC structures allow for complex shapes to be formed by a relatively simple manufacturing process. In addition, the non-horizontal stacking of CMC members substantially reduces or prevents leakage between adjacent laminates relative to conventional horizontally stacked CMC structures. In this way, the components described herein provide improved cooling of the airfoil (thereby allowing for higher temperature use). Still further, the fiber orientation(s) of the CMC material may be optimized in ways not previously possible previously to allow for CMC components having different strength and temperature resistance profiles than previously offered, thereby also enabling higher temperature use. Higher temperature use may, in turn, lead to increased gas turbine operation efficiency.

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

What we claim is:

1. A component (110) comprising:

a plurality of ceramic matrix composite members (111 ) stacked on one another to define an airfoil (116), the airfoil (116) having a radial axis (130) extending therethrough and a horizontal axis (132) perpendicular to the radial axis (132), the ceramic matrix composite members (111 ) each having a degree of curvature (119) and stacked on one another non-parallel to the horizontal axis (132).

2. The component (110) of claim 1 , wherein the degree of curvature (119) follows a camber line (134) of the airfoil.

3. The component (110) of claim 1 , wherein the members (111 ) are stacked on one another at an angle from 10 to 90 degrees relative to the horizontal axis (132).

4. The component (110) of claim 1 , wherein the ceramic matrix composite members (111 ) are stacked on one another parallel to the radial axis (130).

5. The component (110) of claim 1 , wherein the ceramic matrix composite members (111 ) comprise plies (115) of a ceramic matrix composite material.

6. The component (110) of claim 1 , wherein the ceramic matrix composite members (111 ) comprises laminates (112) of a ceramic matrix composite material.

7. The component (110) of claim 1 , further comprising a plurality of cooling passages (142) extending through a body (117) of the airfoil (116).

8. The component (110) of claim 1 , wherein the ceramic matrix composite material comprises an oxide-oxide material.

9. The component (110) of claim 1 , further comprising a thermal barrier coating (152) on an outer surface of the airfoil (116). 10. The component (110) of claims 1 to 9, wherein the component

(110) comprises a rotating component of a gas turbine engine (2).

11. The component of claims 1 to 9, wherein the component (110) comprises a stationary component of a gas turbine engine (2).

12. A process for manufacturing a gas turbine component (110) comprising:

stacking a plurality of ceramic matrix composite members (111 ) non- horizontally on one another to define a precursor block (150) comprising the ceramic matrix composite members (111 ); and

heating the precursor block (150) for a duration and at a temperature effective to sinter the precursor block (150).

13. The process of claim 12, wherein the precursor block (150) defines an airfoil (116).

14. The process of claim 12, further comprising shaping the precursor block (150) to form an airfoil (1 16), the airfoil (1 16) having a radial axis (130) extending therethrough and a horizontal axis (132) perpendicular to the radial axis (130), the ceramic matrix composite members (1 1 1 ) each having a degree of curvature (1 19) and stacked on one another non-parallel to the horizontal axis (132).

15. The process of claim 12, wherein the degree of curvature (1 19) follows a camber line (134) of the airfoil (1 16).

16. The process of claim 12, wherein the ceramic matrix composite members (1 1 1 ) are stacked perpendicular to the horizontal axis (132).

17. The process of claim 12, wherein the ceramic matrix composite members (1 1 ) comprise plies (1 15) of a ceramic matrix composite material.

18. The process of claim 12, wherein the ceramic matrix composite members (1 1 1 ) comprises laminates (1 12) of a ceramic matrix composite material.

19. The process of claim 12, further comprising depositing a thermal barrier coating (152) on outer wall of the airfoil (1 16).

20. The process of claims 12 to 19, wherein the component (1 10) formed comprises a component of a gas turbine engine (2).