FIELD

The present invention relates to ceramic matrix composite (CMC) components, and more particularly to three dimensional (3D) reinforced CMC components which provide for greater thermal resistance and pressure containment relative to

conventional metal, CMC, or hybrid CMC/metal structures.

BACKGROUND

It is well understood that an ability of a system’s components to withstand certain operating temperatures may contribute significantly to the lifetime and efficiency of the system. By way of example, gas turbines include components that are repeatedly exposed to hot gases resulting from combustion of a fuel source. Typically, the gas turbine comprises 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 high 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 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 such 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 or ceramic matrix reinforced with ceramic fibers.

While CMC materials provide excellent thermal protection properties relative to superalloy materials, the mechanical strength of CMC materials is still notably less than that of corresponding high temperature superalloy materials. In particular, laminated (2D) CMC materials have low interlaminar strength, and thus are prone to delamination or other structural damage during high temperature operation. In addition, due to their low thermal conductivity and heat transfer coefficient, CMC materials are difficult to cool. By way of example, FIG. 1 illustrates an airfoil 200 formed from a CMC material having an outer wall 202 defining a large cavity 204 through which a cooling air may flow. During high temperature operation, the flow of cooling air through the cavity 204 results in a high pressure differential between the interior of the cavity and an outside of the airfoil 200. Unfortunately, the high internal pressure will likely result in damage to the CMC outer wall 202 or outright splitting of the airfoil 200 at the trailing edge 206. In addition, the flow of air through cavity 204 will have little effect on cooling the CMC material closest to an exterior 208 of the airfoil 200 since CMC materials have a low thermal conductivity.

Further, introducing internal wall cooling channels in two dimensional (2D) laminated CMC components to cool the component has been demonstrated (see US 6,746,755 and US 8,257,809). Flowever, even with some cooling effect, such components have limited internal pressure containment capability due to the low interlaminar strength of 2D CMCs. For large land-based turbine front stage airfoils, the front stage airfoils should be able to withstand an airfoil internal pressurization and/or pressurization of internal wall cooling channels as high as 10 bar or more. Thus, there is a need for CMC structures with greater robustness for such internal pressurization.

To mitigate the internal pressure, airfoils 220 have further been manufactured with one or more internal ribs 210 which span between the pressure side 212 and the suction side 214 of the airfoil 200 as shown in FIG. 2. While the ribs 210 assist in reducing the stresses associated with the high internal pressure, particularly at the leading edge 216 and the trailing edge 206, the T-joints (where the ribs 210 meet the outer wall 202) cannot withstand the high pressures and are prone to failure. These issues reduce the temperatures within which CMC components may be safely operated. In the context of a turbine engine, reducing the operating temperature of the turbine in turn further reduces turbine efficiency. Accordingly, solutions are still needed for the internal pressurization and cooling issues associated with conventional CMC-containing airfoils.

SUMMARY

To address these needs and others, there are disclosed reinforced ceramic matrix composite (CMC) components which provide novel and inventive reinforced structures relative to known CMC components. In accordance with an aspect, there is provided a component comprising an airfoil, the airfoil comprising fiber material hosted with a ceramic matrix, the fiber material woven in successive layers about spanwise extending reinforcement members to define the airfoil.

In one aspect, there is disclosed a component comprising an airfoil comprising a body of a ceramic matrix composite (CMC) material. The body is defined between a leading edge and a trailing edge, and between a concave portion defining a pressure side and a convex portion defining a suction side, the pressure side and suction side extending between the leading edge and trailing edge. A plurality of reinforcement members extend through the airfoil in a spanwise direction between a root and a tip of the airfoil. The reinforcement members are oriented about a perimeter of the body of the airfoil to generally define a shape of the airfoil. The CMC material comprises a fiber material hosted within a ceramic matrix. The fiber material is woven about the reinforcement members in an axial direction of the airfoil and in successive layers within the ceramic matrix.

In accordance with yet another aspect, there is provided a method of forming a component having an airfoil comprising a body of a ceramic matrix composite material. The body is defined between a leading edge and a trailing edge, and between a concave portion defining a pressure side and a convex portion defining a suction side, the pressure side and suction side extending between the leading edge and trailing edge. The method comprises:

positioning a plurality of spanwise extending reinforcement members to define a general outline of a shape of the airfoil;

weaving a fiber material about the spanwise extending reinforcement members in an axial direction of the airfoil and in successive layers to define the body of the airfoil, wherein the fiber material comprises an amount of a ceramic material effective to produce the ceramic matrix composite material upon sintering of the fiber material and the ceramic material; and

sintering the fiber material with ceramic material woven about the reinforcement members at a temperature and for a duration effective to produce the ceramic matrix composite material and form the body of the airfoil.

BRIEF DESCRIPTION OF THE DRAWINGS

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

FIG. 1 illustrates an embodiment of a PRIOR ART airfoil.

FIG. 2 illustrates another embodiment of a PRIOR ART airfoil.

FIG. 3 illustrates a schematic of a gas turbine engine which incorporates a reinforced woven CMC component in accordance with an aspect of the present invention.

FIG. 4 illustrates a reinforced woven CMC component in accordance with an aspect of the present invention.

FIG. 5 illustrates a cross section of the airfoil in FIG. 4 illustrating fiber material woven about reinforcement members in accordance with an aspect of the present invention.

FIG. 6 illustrates a reinforcement member in accordance with an aspect of the present invention.

FIG. 7 illustrates a reinforcement member in accordance with another aspect of the present invention. FIGS. 8-9 illustrate illustrates a reinforcement member with a fugitive material and with the same removed to form a cooling channel in accordance with yet another aspect of the present invention.

FIG. 10 illustrates a cross section of an airfoil having fiber material woven about reinforcement members and defining ribs in accordance with an aspect of the present invention.

FIG. 1 1 illustrates a cross section of an airfoil having fiber material woven about reinforcement members with a greater concentration of its reinforcement members at the trailing edge in accordance with an aspect of the present invention.

FIG. 12 illustrates a cross section of an airfoil having fiber material woven about every couple reinforcement members in accordance with an aspect of the present invention.

FIG. 13 illustrates an outer surface of the airfoil as comprising a plurality of depressions due to the woven fiber material in accordance with an aspect of the present invention.

FIG. 14 illustrates a schematic of a method for making a reinforced woven CMC component in accordance with an aspect of the present invention.

FIG. 15 illustrates a substrate comprising a plurality of reinforcement members extending spanwise from the substrate in accordance with an aspect of the present invention.

FIG. 16 illustrates fiber material being woven about the reinforcement members shown in FIG. 15 in accordance with an aspect of the present invention.

DETAILED DESCRIPTION

Referring now too the figures, FIG. 3 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 airfoils 16, 18 extend spanwise along a radial direction of the axis 12 of the gas turbine engine 2. 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 now to FIG. 4, there is shown a reinforced component 30 in

accordance with an aspect of the present invention which may comprise a gas turbine component as was illustrated in FIG. 3. In the embodiment of FIG. 4, the component 30 comprises a vane 18 having an elongated airfoil 32, the airfoil 32 having a body 34 which extends in a spanwise (radial) direction (R). The body 34 is defined between a leading edge 54 and a trailing edge 56, and further includes an outer wall 36. The outer wall 36 may have a generally concave-shaped portion 38 defining a pressure side 40 and a generally convex shaped portion (opposite side) 42 defining the suction side 44. The airfoil 32 is disposed between an outer platform 46 at a first end 48 of the vane 18 and an inner platform 50 at a second end 52 of the vane 18. Although a vane 18 is shown, it is appreciated that the component 30 is not limited to a vane 18, but may include any component for high temperature use, such as another component of a turbine engine 2 shown in FIG. 3, e.g., a turbine blade 16.

Referring to FIG. 5, there is shown a cross-section taken at line A-A of the component 30 in FIG.4. As shown, the airfoil 32 comprises a plurality of spanwise extending reinforcement members 55 extending through the body 34 of the airfoil 32 in a spanwise (radial) direction (R). The reinforcement members 55 extend between a root 58 and a tip 60 of the airfoil 32 (FIG. 4). In addition, the reinforcement members 55 are oriented about a perimeter of the body 34 of the airfoil 32 such that when fiber material 62 is woven around the reinforcement members 55 on a layer by layer basis, the body 34 of the airfoil 32 is formed with an airfoil shape. As such, the reinforcement members 55 may be oriented in a general shape or outline of a desired shape or profile (e.g., airfoil shape) of the component 30.

As is also shown in FIG. 5, the airfoil 32 comprises the fiber material 62 woven about the reinforcement members 55 on a layer by layer basis to build the airfoil 32. The fiber material 62 is hosted within a ceramic matrix material 63 to form the ceramic matrix composite (CMC) material 65. In certain embodiments, the fiber material 62 (e.g., a continuous fiber bundle) is woven continuously about the reinforcement members 55 from the root 58 to the tip 60 of the airfoil 32. In this way, the fiber material

62 is provided as a single piece or unit, cutting of the fiber material 32 is eliminated, and the mechanical strength of the CMC material 65 may be uniform throughout the component 30 with no weak points.

In an aspect, the fiber material 62 is woven about the reinforcement members 55 in an axial direction of the airfoil 32, and in successive layers within the ceramic matrix

63 to build the component 30. In one embodiment, the axial direction 68 extends in a direction between the leading edge 54 and the trailing edge 56 or vice-versa. In other embodiments, the axial direction 68 may further or instead extend in a direction between the pressure side 40 and the suction side 44. The axial direction may extend at any angle tangential to a spanwise direction (R) extending through the airfoil 32, and may also comprise a component perpendicular to the spanwise direction (R).

The reinforcement members 55 may comprise any material having a rigidity effective to provide at least a degree of reinforcement to the body 34 of the airfoil 32 in the axial and/or spanwise direction against internal pressure forces (see arrows in FIGS. 1 -2) and/or compressive forces. In certain embodiments, as shown in FIG. 6, the reinforcement members 55 comprise a solid material 72 - meaning that any given cross section normal to a spanwise (radial) axis thereof is solid. In certain embodiments, the reinforcement members 55 comprise a ceramic material, a ceramic matrix, or a ceramic matrix composite material. In other embodiments, the reinforcement members may comprise a carbon material. In accordance with another aspect and as shown in FIG. 7, the reinforcement members 55 comprise a body 74 which defines a bore 76 extending therethrough in the spanwise direction (R). The bore 76, in turn, defines a cooling channel 78 for each respective reinforcement member 55 in the airfoil 32. The reinforcement members 55 comprising a cooling channel 78 may similarly be formed of a carbon material, a ceramic material, a ceramic matrix material, or a ceramic matrix composite material as with a solid reinforcement member 5.

In certain embodiments, the reinforcement members 55 themselves may comprise a fiber material 62. The fiber material 62 may be in any suitable form that provides the reinforcement members 55 with a degree of rigidity or reinforcement as discussed above, such as in the form of fiber bundles, braids, ropes, or the like. In an embodiment, the reinforcement members 55 comprise at least one of a braided ceramic rope or a hollow braided ceramic rope having a bore 76 (cooling channel 78) extending therethrough in a lengthwise or spanwise dimension of the fiber material 62. In certain embodiments, as shown in FIG. 8, the reinforcement members 55 may be provided with a bore 76 which, prior to a heat treatment (firing), is filled with a fugitive material 80. In this way, the reinforcement member 55 comprises an added degree of structural strength to the body 32 before firing of the same. The fugitive material 80 may comprise a wax material or a low melting temperature polymer (at least having a lower melting point than the body 74), for example. In any case, the fugitive material 80 may be melted and removed from the component 30 during heat treatment to leave behind a cooling channel 78 in selected reinforcement members 55 as shown in FIG. 9.

The reinforcement members 55 may be distributed in any suitable arrangement throughout a body 34 of the airfoil 32. In addition, any suitable number or concentration of reinforcement members 55 may be provided. For example, in an embodiment, the reinforcement members 55 may be provided in a single row to define a perimeter 90 of the airfoil as was shown in FIG. 4. In other embodiments, multiple rows of

reinforcement members 55 may be provided about the perimeter 90 of the airfoil 32. In certain embodiments, as shown in FIG. 10, the reinforcement members 55 may also be utilized to define ribs 82 in the airfoil 32 defining one or more (cooling) cavities 84 in the airfoil 32. It is appreciated that any given number of the reinforcement members 55 may be provided in any desired location to provide a degree of desired degree of reinforcement to the airfoil 32. For example, the reinforcement members 55 may be provided with any desired spacing between adjacent members 55. In certain

embodiments, the spacing between members 55 aids in defining a density of the members 55 in a given region of the airfoil 32. In particular embodiments, the density of the reinforcement members 55 is greater (e.g., more reinforcement members 55 per a given unit of area) at the trailing edge 54 of the airfoil 32 relative to a remaining portion of the airfoil 32 as shown in FIG. 11.

The ceramic matrix composite (CMC) material 65 comprises a fiber material 62 hosted with a ceramic matrix material 63. The fiber material 62 may comprise an oxide material, a non-oxide material, or a combination thereof, such as alumina, mullite, aluminosilicate, ytrria alumina garnet, silicon carbide, silicon nitride, silicon carbonitride, and the like, and combinations thereof. It is appreciated that the CMC material 65 may combine a matrix composition with a reinforcing phase (fiber material 62) of a different composition (such as mullite/silica), or may be of the same composition

(alumina/alumina or silicon carbide/silicon carbide). In an embodiment, the CMC material 65 comprises an oxide-oxide (Ox-Ox) CMC material having oxide fibers disposed within an oxide matrix. The fibers of the fiber material 62 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. The (host) ceramic matrix material 63 may further contain whiskers, platelets, particulates, or fugitives, or the like for additional reinforcement.

The fiber material 62 can be woven about the reinforcement members 55 by any suitable method on a layer by layer basis to build the airfoil 32 in the spanwise direction (R) (until complete). In accordance with one aspect and referring again to FIG. 5, the fiber material 62 may be woven about the members 55 such that at least a portion of the fiber material 62 travels over selected reinforcement members 55 in a first pass and under the (same) selected reinforcement members in a second pass. This may be done repeatedly so as to build the body 34 of the airfoil 32. In certain embodiments and as is shown in FIG. 5, the fiber material 62 travels over each reinforcement member 55 in the first pass 86 and then back under the same reinforcement member 55 in the second pass 88.

In other embodiments, as shown in FIG. 12, the fiber material 62 is woven such that it travels over two selected reinforcement members in the first pass 86 and under the (same) selected reinforcement members in a second pass 88. In certain

embodiments, the same pattern is repeated over and over to build the body 34 of the airfoil 32 in the spanwise direction (R) such that the fiber material 62 travels over and under given reinforcement members 55 in the spanwise direction (R). In certain embodiments, however, the layers of fiber material 62 may be staggered such that a first layer of the fiber material 62 travels over different reinforcement members relative 55 to a second layer of the fiber material 62, wherein the second layer is disposed above or below the first layer in the spanwise direction. This may have the effect of providing differential load bearing capabilities through the structure. For example, the fiber material 62 could be woven over and under each reinforcement member toward a root 58 of the airfoil 32 (FIG. 5), and thereafter woven over and under every two reinforcement members towards a tip 60 of the airfoil 32 (FIG. 12). It is understood that various weaving patterns of the fiber material 62 may be achieved in this manner by the size and placement of the reinforcing members 55, the size and shape (e.g., served or unserved) of the fiber material 62 (e.g., weaving fiber bundle), the axial tension applied to the fiber material 62 (e.g., weaving yarn), and the radial packing pressure (battening) applied.

In addition, as mentioned previously, the reinforcement members 55 may be located at a position effective to form one or more ribs 82. Since the ribs 82 are integral with the outer perimeter 90 of the body 32 of the airfoil 32, the ribs 82 formed as described herein are significantly less likely to fail than ribs in conventional structures - even upon being subjected to the high internal pressures expected within cavities 84 during high temperature operation of the component 30 with airfoil 32. Conventional CMC structures, particularly though comprising plies, are particularly weak at a T-joint at the junction between their ribs and the outer perimeter of the airfoil as the ribs are necessarily thin and are not strongly anchored to the outer perimeter body portion of the airfoil. In contrast, the ribs 82 described herein are integrally formed with the outer perimeter 90 via the fiber material 62. In an embodiment and as shown in FIG. 10, the airfoil 32 comprises a rib 82 defining a cavity 84 between the outer perimeter 90 of the body 34 of the airfoil 32 and the rib 82, wherein the fiber material 62 extends from the pressure side 40 to the suction side 44 of the airfoil 32 through the rib 82.

In accordance with another aspect, the trailing edge of the airfoil 32 may also be susceptible to splitting as a result of the pressure created within the internal cavity or cavities of the airfoil. In accordance with another aspect of the present invention, the fiber material 62 may be woven at the trailing edge 56 such that its mechanical strength is increased relative to a remainder of the airfoil 32 at the trailing edge. By way of example and as shown in FIG. 11 , the airfoil 32 may comprise a greater number of reinforcement members 55 per unit area at the trailing edge 56. In other embodiments, larger or mechanically stronger reinforcement members 55 may be utilized at the trailing edge 56 relative to a remainder of the airfoil 32.

To form the ceramic matrix composite (CMC) material 65, the fiber material 62 is impregnated with an effective amount of a ceramic or ceramic precursor material (ceramic material 63) as described herein. The fiber material 62 may be impregnated with the ceramic material 63 by any suitable method. In certain embodiments, the fiber material 62 is pre-impregnated with a ceramic material 63 prior to the weaving about the fiber material 62 about the reinforcement members 55. In other embodiments, the fiber material 62 is woven about the reinforcement members 55 in their desired position, impregnated with a ceramic or ceramic precursor material, and sintered to form the desired CMC material 65. The ceramic material 63 itself may comprise any suitable oxide, non-oxide material, or combinations thereof, such as one or more of alumina, mullite, aluminosilicate, yttria alumina garnet, silicon carbide, silicon nitride, silicon carbonitride, and the like, or a precursor thereof. In particular embodiments, the ceramic material 63 comprises an oxide material. In any case, once the fiber material 62 is impregnated with the ceramic material, the impregnated fiber material 62 may be subjected to a sintering process in order to provide the final ceramic matrix composite material 65. As is known in the art, the sintering process may be done at a temperature and for a duration effective to ensure flow of the ceramic material 63 into or about the fiber material 62 and suitable drying to form the desired ceramic matrix composite material 65. In an embodiment, the sintering is done at a temperature of from 500 to 1300° C, isocratic or with a gradient, for a duration of from 1 to 24 hours.

In accordance with another aspect, the reinforced woven structure of the present invention substantially increases a structural strength of the associated airfoil 32. In certain embodiments, the one or more cavities 84 defined within an interior of the airfoil 32 may be provided with a metal support column (not shown) that also extends in the spanwise direction (R) through the one or more cavities 84 of the airfoil 32. The metal material of the column may comprise a suitable 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.

In accordance with yet another aspect, in certain embodiments, the final CMC material 65 comprises a non-smooth three-dimensional surface as a result of the woven fiber material 62 about the reinforcement members 55. In certain embodiments and as shown in FIG. 13, the depressions 91 formed in the CMC material 65 provides the airfoil 32 with an improved surface for the attachment of a thermal barrier coating (TBC) thereto in order to further improve the thermal resistance of the airfoil 32 if desired. A TBC is thus an optional coating to the embodiments described herein to provide an added degree of thermal protection, e.g., maintain lower exposure temperatures, to the underlying CMC material 65. The TBC may comprise any suitable material which provides an increased temperature resistance to the CMC material 65 when applied thereover. In an embodiment, the TBC comprises a stabilized zirconia material. For example, the TBC 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 may comprise a magnesia stabilized zirconia, ceria stabilized zirconia, aluminum silicate, or the like.

In another embodiment, the TBC may comprise a pyrochlore structure. In an embodiment, the pyrochlore structure has the empirical formula A2B2O7 or in general terms AvBxOz 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 AvBxOz 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., Gd _v (Hf _x Zr _y )Oz), x + y = 2. In other embodiments, B = Zr or Hf individually, and the pyrochlore structure may comprise one of gadolinium zirconate (Gd ₂ Zr ₂ 07) or gadolinium hafnate (Gd ₂ Hf ₂ 07). In still other embodiments, the TBC 92 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 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 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 may further comprise a degree of porosity suitable for the desired application.

In accordance with another aspect of the present invention, there is shown in FIG. 14 and described below a method 100 for making a airfoil 32 as described herein. The elements in the method 100 have been described in detail above and are not restated herein in the interest of brevity. In an embodiment, the method 100 includes step 102 of positioning a plurality of the reinforcement members 55 to define a general outline of a shape of an airfoil 32. In an embodiment and as shown in FIG. 15, the reinforcement members 55 may be threaded into corresponding apertures 105 in a substrate 104 or otherwise affixed to the substrate 104 upon which the airfoil 32 will be built. In certain embodiments, the substrate 104 comprises a platform 106 as typically utilized in a vane 18 (FIG. 4).

Thereafter, as shown in FIG. 16, the method 100 further includes step 105 of weaving the fiber material 62 about spanwise extending reinforcement members 55 in the axial direction 68 and in successive layers to define the body of the airfoil 32. This is done in any suitable pattern and continues until an amount of the fiber material 62 is woven about the reinforcement members 55 sufficient to build an airfoil 32 of a desired shape and size. In certain embodiment, the fiber material 62 comprises more than one continuous length of a braided or non-braided fiber materials. Thus, in certain embodiments, a first length the fiber material 62 may be wound around the

reinforcement members 55 and then cut and a new (second) length started. In other embodiments, the fiber material 62 is continuous and comprises a single length of fibers (e.g., straight, braided, bundled, etc.) continuously woven about the reinforcement members 55 layer by layer to build the body 34 of the airfoil 32 in the spanwise direction (R). The continuous weaving of a single fiber strand provides the airfoil 32 with maximum reinforcement properties.

The fiber material 62 comprises an amount of a ceramic material 63 effective to produce the ceramic matrix composite material 65 upon sintering of the fiber material 62 and the ceramic material 63. As discussed above, the fiber material 62 may be provided with the ceramic material 63 by any suitable process. For example, the fiber material 62 may be pre-impregnated with the ceramic material 63 prior to the weaving the fiber material 62 about the reinforcement members 55. In still other embodiments, the fiber material 62 may be woven about the reinforcement members 55. Thereafter, the woven fiber material 62 may be impregnated with a ceramic material 63 as described herein, which may comprise a composition comprising ceramic particles dispersed within a suitable medium.

Once impregnated, the method further includes step 112 of sintering the fiber material 62 with ceramic material 63 woven about the reinforcement members 55 at a temperature and for a duration effective to produce the ceramic matrix composite material 65 and form the body 34 of the airfoil 32. In an embodiment, the sintering is done at a temperature of from 500 to 1300° C, isocratic or with a gradient, for a duration of from 1 to 24 hours. In certain embodiment, as described above, the reinforcement members may further define cooling channels 78 within the reinforcement members 55 - either by the presence of a bore extending therethrough in the spanwise direction (R). In other embodiments, a fugitive material 80 is included within a bore of the

reinforcement members 55, which may be removed, e.g., melted away, during the sintering step 1 12.

It is known that CMC materials, particularly those formed from a plurality of plies of ceramic material, include weak interlaminar strength. The components and processes described herein produce turbine component having a three-dimensional woven fiber material that significantly increases a mechanical strength of the

component. In addition, the strengthened CMC component significantly cooling needs for the component relative to conventional CMC, superalloy, and/or hybrid (CMC and superalloy) components due to its increased thermal resistance.

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.