Patent Description:
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.

Airfoils in the turbine section are typically formed of a superalloy and may include thermal barrier coatings to extend temperature resistance. Ceramic matrix composite ("CMC") materials are also being considered for airfoils. Among other properties, CMCs have high temperature resistance and oxidation resistance.

<CIT> describes a ceramic matrix composite material (CMC) vane for a gas turbine engine wherein the airfoil member and the platform member are formed separately and are then bonded together to form an integral vane component. The airfoil member and the platform member may be bonded together by an adhesive after being fully cured. Alternatively, respective joint surfaces of the green body state airfoil member and platform member may be co-fired together to form a sinter bond. A mechanical fastener and/or a CMC doubler may be utilized to reinforce the bonded joint. A matrix infiltration process may be used to create or to further strengthen the bond.

According to a first aspect of the present invention as set forth in claim <NUM>, there is provided an airfoil assembly for a gas turbine engine comprising at least one inner ceramic matrix composite ply which defines an internal cavity of an airfoil piece and first and second collar projections. The at least one inner ceramic matrix composite ply is continuous through the airfoil piece and first and second collar projections. First and second platforms are at first and second ends of the airfoil piece. The first and second collar projections extend radially past the first and second platforms, respectively. The first collar projection has a length in the radial dimension that is longer than a length of the second collar projection in the radial dimension.

In an embodiment of the above embodiment, the airfoil assembly includes a first outer ceramic matrix composite ply. The first outer ceramic matrix composite ply surrounds the inner ceramic matrix composite ply at the first collar projection and defines a radially outer surface of the first platform. The airfoil piece also includes a third outer ceramic matrix composite ply. The third outer ceramic matrix composite ply surrounds the inner ceramic matrix composite ply at the second collar projection and defines a radially inner surface of the second platform.

In an embodiment of any of the above embodiments, a second outer ceramic matrix composite ply defines a radially inner surface of the first platform, a radially outer surface of the second platform and an outermost surface of the airfoil section, and surrounds the inner ceramic matrix composite ply.

In an embodiment of any of the above embodiments, a middle ceramic matrix composite ply surrounds substantially the entire radial length of the inner ceramic matrix composite ply.

In an embodiment of any of the above embodiments, the at least one inner ceramic matrix composite ply includes a first inner ceramic matrix composite ply which defines a first cavity and a second inner ceramic matrix composite ply which defines a second cavity.

In an embodiment of any of the above embodiments, a middle ceramic matrix composite ply surrounds the first and second inner ceramic matrix composite plies to define the airfoil piece.

In embodiment of any of the above embodiments, the first inner ceramic matrix composite ply comprises ceramic-based fibers in a ceramic-based matrix. At least some of the fibers are continuous through the first and second collar projections and the airfoil piece.

In embodiment of any of the above embodiments, there is provided a vane for a gas turbine engine comprising an airfoil assembly. A spar piece includes a hollow spar situated in the airfoil cavity and a spar platform. The spar platform includes a pocket. The first collar projection is situated in the pocket.

In an embodiment of the above embodiments, the first collar projection is a radially outer collar projection and the second collar projection is a radially inner collar projection.

In an embodiment of any of the above embodiments, the vane includes an outer ply, the outer ply is a second outer ceramic matrix composite ply and the second outer ceramic matrix composite ply defines a radially inner side of the first platform, a radially outer side of the second platform and an outermost surface of the airfoil section. A first outer ceramic matrix composite ply defines a radially outer side of the first platform and a third outer ply defines a radially inner side of the second platform.

In an embodiment of any of the above embodiments, the first and third outer plies surround a middle ply in the first and second collar projections, respectively.

In an embodiment of any of the above embodiments, the airfoil cavity is a first airfoil cavity and the inner ply is a first inner ply, and the airfoil further includes a second airfoil cavity which is defined by a second inner ply. The middle ply surrounds the first and second inner plies.

In an embodiment of any of the above embodiments, the spar piece is configured to transfer structural loads from the airfoil piece to a support structure.

According to a second aspect of the present invention as set forth in claim <NUM>, there is provided a method of assembling a vane comprising inserting a spar piece into a cavity of a hollow airfoil section of an airfoil piece. The cavity is defined by an inner ceramic matrix composite ply. A middle ceramic matrix composite ply surrounds the inner ply, and an outer ceramic matrix composite ply surrounds the middle ply. The outer ply defines first and second platforms. The inner and middle plies extend radially past the first and second platforms to form first and second collar projections. The first collar projection has a length in the radial dimension that is longer than a length of the second collar projection in the radial dimension.

In an embodiment of the above embodiments, inserting includes aligning a pocket in a spar platform of the spar piece with the first collar projection.

In an embodiment of any of the above embodiments, the outer ply is a second outer ceramic matrix composite ply and the second outer ceramic matrix composite ply defines a radially inner side of the first platform, a radially outer side of the second platform and an outermost surface of the airfoil section. A first outer ceramic matrix composite ply defines a radially outer side of the first platform and a third outer ceramic matrix composite ply defines a radially inner side of the second platform.

In an embodiment of any of the above embodiments, the spar piece is metallic.

Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including but not limited to three-spool architectures.

Terms such as "axial," "radial," "circumferential," and variations of these terms are made with reference to the engine central axis A.

In a further example, the engine <NUM> bypass ratio is greater than about <NUM>:<NUM>, with an example embodiment being greater than about <NUM>:<NUM>, the geared architecture <NUM> is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM> and the low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>:<NUM>. In one disclosed embodiment, the engine <NUM> bypass ratio is greater than about <NUM>:<NUM>, the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>:<NUM>. The low pressure turbine <NUM> pressure ratio is pressure measured prior to the inlet of low pressure turbine <NUM> as related to the pressure at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including but not limited to direct drive turbofans.

The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about <NUM>:<NUM>.

<FIG> illustrates a representative vane <NUM> from the turbine section <NUM> of the engine <NUM>, although the examples herein may also be applied to vanes in the compressor section <NUM>. A plurality of vanes <NUM> are situated in a circumferential row about the engine central axis A. <FIG> show a detail view of a radially outer end of the vane <NUM>, although it is to be appreciated that modified examples include the radially inner end. <FIG> shows a cross-sectional view of the radially outer end of the vane <NUM> taken along the section line A-A in <FIG>.

The vane <NUM> is comprised of an airfoil piece <NUM> and a spar piece <NUM> (<FIG>). The airfoil piece <NUM> includes several sections, including first (radially outer) and second (radially inner) platforms <NUM>/<NUM> and a hollow airfoil section <NUM> that joins the first and second platforms <NUM>/<NUM>. The airfoil section <NUM> includes at least one cavity <NUM>. In this example, there are three cavities <NUM> though in other examples more or less cavities <NUM> could be used. The airfoil section <NUM> extends beyond the first platform <NUM> to form a collar <NUM> that projects radially from the first platform <NUM>, i.e. the collar projection <NUM> is an extension of the airfoil section from the first platform <NUM> and thus continues the shape profile of the airfoil section. As shown in <FIG>, the inner platform <NUM> can also include a collar projection <NUM>. The terminology "first" and "second" as used herein is to differentiate that there are two architecturally distinct components or features. It is to be further understood that the terms "first" and "second" are interchangeable in the embodiments herein in that a first component or feature could alternatively be termed as the second component or feature, and vice versa.

The airfoil piece <NUM> may be formed of a metallic material, such as a nickel- or cobalt-based superalloy, but more typically will be formed of a ceramic. The ceramic may be a ceramic matrix composite ("CMC"). Example ceramic materials may include, but are not limited to, silicon-containing ceramics. The silicon-containing ceramic may be, but is not limited to, silicon carbide (SiC) or silicon nitride (Si<NUM>N<NUM>). An example CMC may be a SiC/SiC CMC in which SiC fibers are disposed within a SiC matrix. The CMC may be comprised of fiber plies that are arranged in a stacked configuration and formed to the desired geometry of the airfoil piece <NUM>. For instance, the fiber plies may be layers or tapes that are laid-up one on top of the other to form the stacked configuration. The fiber plies may be woven, unidirectional, knitted, or braided, for example. In one example, at least a portion of the fiber plies may be continuous through the first platform <NUM>, the airfoil section <NUM>, and the second platform <NUM>. In this regard, the airfoil piece <NUM> may be continuous in that at least some of the fiber plies are uninterrupted through the first platform <NUM>, the airfoil section <NUM>, and the second platform <NUM>, as discussed in more detail below. In alternate examples, the airfoil piece <NUM> may be discontinuous such that the first platform <NUM>, the airfoil section <NUM>, and/or the second platform <NUM> are individual sub-pieces that are attached to the other sections of the airfoil piece <NUM> in a joint.

The spar piece <NUM> defines a spar platform <NUM> and a (hollow) spar <NUM> that extends from the spar platform <NUM> into the hollow airfoil section <NUM>. For example, the spar piece <NUM> is formed of a metallic material, such as a nickel- or cobalt-based superalloy, and is a single, monolithic piece. The spar piece <NUM> includes a radial pocket <NUM> which receives the collar projection <NUM>. The spar piece <NUM> is configured to connect to a support structure (not shown) of the engine <NUM>. The spar piece <NUM> bears structural loads from the airfoil piece <NUM> during operation of the engine <NUM>. In particular, the airfoil piece <NUM> transfers loads directly to the spar piece <NUM> via the interaction of the collar projection <NUM> and the pocket <NUM> in the spar platform <NUM>. The platform <NUM>/<NUM> and collar projections <NUM> also act as a heat shield for the spar platform <NUM>.

As best shown in <FIG>, the airfoil piece <NUM> includes multiple plies of CMC material, as discussed above. The plies are formed and joined together by any known manner. Furthermore, the plies can be formed of the same or different materials. It should be understood that the plies discussed herein can also be sets of multiple plies, in other examples.

Inner plies <NUM> define each cavity <NUM>. The inner plies <NUM> can be shaped to define the cavity <NUM> by being formed on a mandrel, as would be known in the art. As shown, the inner plies <NUM> are continuous through the entire radial length of the airfoil piece <NUM>, including the collar projections <NUM> and the airfoil section <NUM>. A middle ply <NUM> (<FIG>) surrounds substantially the entire length of the inner ply <NUM> to define the outer shape of the airfoil section <NUM>, join the individual inner plies <NUM> together, and provide structural integrity to the airfoil piece <NUM>. Like the inner plies <NUM>, the middle ply <NUM> is continuous through the entire radial length of the airfoil piece <NUM>, including the collar projections <NUM> and the airfoil section <NUM>. The continuous nature of the inner plies <NUM> and middle ply <NUM> improves the strength of the airfoil piece <NUM> so that it can withstand and transfer loads directly to the spar piece <NUM> as discussed above. As discussed above, the plies include ceramic-based fibers in a ceramic-based matrix. In one example, the continuous plies include at least some fibers that are continuous through the entire radial length of the airfoil piece <NUM>, including the collar projection <NUM> and the airfoil section <NUM>. That is, the continuous plies include at least some continuous fibers that have a portion disposed in the collar projection <NUM> at one or both platforms <NUM>/<NUM> and a portion disposed in the airfoil section <NUM>. In a particular example, one or both of the inner and middles plies <NUM>/<NUM> comprise ceramic-based fibers arranged in triaxial braids as would be known in the art (though other fiber arrangements are also contemplated). In this example, at least some of the triaxial braids are continuous through the entire radial length of the airfoil piece <NUM>, including the collar projections <NUM> and the airfoil section <NUM>.

Outer plies define the outermost surfaces of the airfoil piece <NUM> and platforms <NUM>/<NUM>. Outer plies includes a first outer ply 104a, which defines a radially outer surface of the platform <NUM> and surrounds the second ply <NUM> in the collar projection <NUM>. A second outer ply 104b defines a radially inner surface of the platform <NUM>, a radially outer surface of the platform <NUM>, and an outermost surface of the airfoil <NUM>, middle ply <NUM>, and third outer ply 104c. It should be understood, however, that more plies could be arranged between the specific plies discussed herein.

The collar projection <NUM> has a length d defined in the radial dimension. The collar projection <NUM> at the radially outer platform <NUM> has a greater length d than the collar projection <NUM> at the radially inner platform <NUM>. The thermal gradients discussed above are generally more pronounced at the radially outer end of the vane <NUM> as compared to the radially inner end. Accordingly, the larger collar projection <NUM> at the radially outer platform <NUM> better mitigates the temperature gradient across the vane <NUM> as discussed above. In a particular example, the length d of the collar projection <NUM> at the outer platform <NUM> is between about <NUM> and <NUM> % of the radial length L of the airfoil piece <NUM> (shown in <FIG> as the length between platforms <NUM>/<NUM>) and the length d of the collar projection <NUM> at the inner platform <NUM> is between about <NUM> and <NUM> % of the radial length L. In another particular example, a ratio of the length d of the collar projection <NUM> at the outer platform <NUM> and the length d of the collar projection <NUM> at the inner platform is about <NUM>:<NUM>.

The vane <NUM> is assembled by inserting the spar piece <NUM> into the airfoil piece <NUM>. The assembly includes aligning the pocket <NUM> with the collar projection <NUM> such that the collar projection <NUM> extends into the pocket <NUM> when the vanes <NUM> are assembled.

Claim 1:
An airfoil assembly for a gas turbine engine (<NUM>), comprising:
at least one inner ceramic matrix composite ply (<NUM>) defining an internal cavity (<NUM>) of an airfoil piece (<NUM>) and first and second collar projections (<NUM>), the at least one inner ceramic matrix composite ply (<NUM>) being continuous through the airfoil piece (<NUM>) and first and second collar projections (<NUM>);
first and second platforms (<NUM>, <NUM>) at first and second ends of the airfoil piece (<NUM>), wherein the first and second collar projections (<NUM>) extend radially past the first and second platforms (<NUM>, <NUM>), respectively; and
characterized in that
the first collar projection (<NUM>) has a length (d) in the radial dimension that is longer than a length (d) of the second collar projection (<NUM>) in the radial dimension.