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 method for securing a nozzle for a turbine. The nozzle includes an airfoil having a suction side and a pressure side connected at a leading edge and a trailing edge such that a cooling cavity is defined within the airfoil, the airfoil extending between an inner band and an outer band. The method includes extending at least one member through the airfoil, and at least one of the inner band and the outer band. The method further includes securing the nozzle assembly in position with at least one fastener such that the at least one member is coupled adjacent to at least one of the inner band and the outer band.

<CIT> describes a hot section component of a gas turbine engine having a covering. The covering includes a protrusion and is attached to the hot section component through a flexible retainer. In one form the covering is made from ceramic matrix composite. The flexible retainer has a closed position and an open position. The retainer secures the protrusion to the hot section component when it engages part of the protrusion when in the closed position.

<CIT> describes a segmented component for use with a gas turbine engine comprising a radially extending gas path portion. The gas path portion is for interacting with gas flow from the gas turbine engine. The gas path portion comprises a forward portion forming a leading edge of a stationary vane, an aft portion forming a trailing edge of the stationary vane, and a plurality of middle portions forming a pressure side and a suction side of the stationary vane. The component is divided into axially aligned segments comprising a forward segment, an aft segment, and a plurality of middle segments disposed between the forward segment and the aft segment. The middle segments comprise radially elongate ceramic matrix composite material plates.

<CIT> describes a turbine assembly for a gas turbine engine with ceramic matrix composite vane disclosing the technical features of the preamble of claim <NUM>.

According to a first aspect of the present invention, there is provided a vane airfoil assembly comprising an airfoil piece comprising a first airfoil platform, a second airfoil platform, and a hollow airfoil section that joins the first airfoil platform and the second airfoil platform. The hollow airfoil section includes a collar projection which extends past the first airfoil platform. The collar projection includes at least one radial tab projecting therefrom. A spar piece defines a spar platform and a spar extends from the spar platform into the hollow airfoil section, the spar piece including a radial pocket defined by first and second opposed faces, the radial pocket configured to receive the collar projection. The hollow airfoil includes first and second cavities separated by a divider, at least one radial tab is at a circumferential location of collar projection that corresponds to a location of the divider.

In an embodiment of the above embodiment, the collar projection has a first radial extent and the at least one radial tab has a second radial extent, and a ratio of the first radial extent to the second radial extent is between about <NUM>:<NUM> and <NUM>:<NUM>.

In an embodiment of any of the above embodiments, at least one radial tab comprises a plurality of radial tabs, and the plurality of radial tabs have a cumulative circumferential extent.

In an embodiment of any of the above embodiments, the cumulative circumferential extent is between about <NUM> and <NUM>% of a circumferential extent of the collar projection.

In an embodiment of any of the above embodiments, the airfoil section and collar projection include at least one continuous ceramic matrix composite ply.

In an embodiment of any of the above embodiments, the hollow airfoil section is configured to receive a metallic spar piece therein.

In an embodiment of any of the above embodiments, the collar projection is a radially outer collar projection.

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 via the collar projection.

In an embodiment of any of the above embodiments, the radial pocket includes at least one mating feature which is configured to mate with the at least one radial tab.

According to a second aspect of the present invention, there is provided a method of assembling a vane comprising inserting a spar piece into a cavity of a hollow vane airfoil section of an airfoil assembly in an embodiment of any of the above embodiments. In an embodiment of the above embodiments, the insertion includes aligning a pocket in a spar platform of the spar piece with the collar projection.

In an embodiment of any of the above embodiments, the pocket includes at least one mating feature which is configured to mate with the at least one radial tab. In an embodiment of any of the above embodiments, the spar piece is metallic and the airfoil piece is ceramic matrix composite.

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 six <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> shows 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> separated by dividers <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 projection <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. In some examples, 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> connects to a support structure in the engine <NUM> (not shown). 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 between collar projection <NUM> and the pocket <NUM> in the spar platform <NUM>. The platform <NUM>/<NUM> and collar projection <NUM> also act as a heat shield for the spar platform <NUM>.

As best shown in <FIG>, the collar projection <NUM> has a radial extent d1 defined from the platform <NUM>. In a particular example, the radial extent d1 of the collar projection <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>).

The vane <NUM> also includes one or more radial tabs <NUM> extending from the collar projection <NUM>. Though in the example shown there are three radial tabs <NUM>, more or less radial tabs could be used. Each of the tabs <NUM> has a radial extent d2. The tabs <NUM> are generally rectangular in shape, and in one example, have rounded outer corners <NUM>.

The tabs <NUM> provide additional mass/surface area for load transfer to the spar piece <NUM> as discussed above. The tabs <NUM> can be located anywhere around the circumference of the airfoil section <NUM>. In one example, the tabs <NUM> are disposed near areas of the airfoil section <NUM> that experience the highest loads. According to the invention, at least one tab <NUM> is at a circumferential location of the collar projection that corresponds to the location of the divider <NUM>. In another particular example, at least one tab <NUM> is located near the trailing edge TE of the airfoil section <NUM>.

The vane <NUM> experiences high heat during operation of the engine <NUM>. The high heat causes thermal expansion of the airfoil piece <NUM> and the spar piece <NUM>. Because the airfoil piece <NUM> is a CMC material and the spar <NUM> is metallic, the airfoil piece <NUM> and spar piece <NUM> thermally expand and contract at different rates. Also, the spar piece <NUM> and/or airfoil section <NUM> receive cooling air in the hollow spar <NUM> such as bleed air from the compressor section <NUM> (<FIG>). The airfoil piece <NUM>, which comprises a CMC material, is prone to relatively higher temperature gradients along its length as compared to the spar piece <NUM>. Accordingly, despite the cooling scheme, there may be a temperature mismatch between the airfoil piece <NUM> and the spar piece <NUM>, which leads to the tendency of heat conduction between the airfoil piece <NUM> and the spar piece <NUM>. The heat conduction contributes to the thermal gradient and tends to increase the gradient. The temperature mismatch is greatest at the radially inner and outer ends of the airfoil piece <NUM>, which receives relatively less cooling air as compared to the airfoil section <NUM>. The tabs <NUM> provide a surface area for load transfer to the spar piece <NUM> as discussed above, but overall minimize the surface area in the radial and circumferential dimensions (e.g., the product of the radial extent d2 and circumferential extent c discussed below) near the radially inner/outer ends of the airfoil piece <NUM> for heat conduction. Therefore, due to the tabs <NUM>, the temperature gradient is overall urged downwards.

As shown in <FIG>, the radial tabs <NUM> have a radial extent d2. In some examples, the radial extent d2 is less than the radial extent d1 of the collar projection <NUM>. In a particular example, the ratio of the radial extents d1:d2 is between about <NUM>:<NUM> and <NUM>:<NUM>. Each of the radial tabs <NUM> also has a circumferential extent c. A cumulative circumferential extent is defined as the sum of the circumferential extent c of each of the tabs <NUM>. In some examples, the cumulative circumferential percent of the radial tabs <NUM> is between about <NUM> and <NUM>% of the circumferential extent of the collar projection <NUM>. The radial tabs <NUM> also have a surface area which is the product of the radial extent d2 and the circumferential extent c. A cumulative surface area is defined as the sum of the surface area of each of the tabs <NUM>.

As discussed above, the airfoil piece <NUM> is formed of CMC plies 100a/100b (<FIG>). Though two plies are schematically shown, more plies could be used. For example, additional plies could be used as an outer wrap around the plies 100a/100b, and/or additional plies could be used to define the cavities <NUM>. The CMC plies 100a/100b are continuous through the airfoil section <NUM> and the collar projection <NUM>. The continuous plies 100a/100b improve the strength of the airfoil section <NUM> and collar projection <NUM>. In turn, the airfoil piece <NUM> withstands and transfers loads directly to the spar piece <NUM> as discussed above. The plies can be layed up, consolidated, and cured as would generally be known in the art. In one example, the tabs <NUM> are machined into the collar projection <NUM> after formation of the airfoil piece <NUM> by any known method. In this example, the collar projection <NUM> is manufactured to a radial extent equal to the sum of the radial extents d1 and d2 (<FIG>). In some examples, the collar projection <NUM> is manufactured with a manufacturing excess. Manufacturing excess is excess material that is then machined down after the airfoil piece <NUM> is formed to provide a desired size and geometry for the airfoil piece, as would be known in the art. Then, the manufacturing excess is removed by machining to form the tabs <NUM>.

In one example, the collar projection <NUM> and tabs <NUM> fill substantially all of the radial extent of the pocket <NUM> in the spar platform <NUM>. In one example, the pocket <NUM> in the spar platform <NUM> includes mating features <NUM> which mate with the tabs <NUM>. The mating features <NUM> locate the tabs <NUM> and collar projection <NUM> with respect to the pocket <NUM>. The mating features <NUM> also are configured to support the tabs <NUM> within the pocket <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> (and optionally the mating features <NUM> with the tabs <NUM>) such that the collar projection <NUM> extends into the pocket <NUM> when the vanes <NUM> are assembled.

Although the different examples are illustrated as having specific components, the examples of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the embodiments in combination with features or components from any of the other embodiments.

Claim 1:
A vane airfoil assembly comprising:
an airfoil piece (<NUM>), comprising:
a first airfoil platform (<NUM>);
a second airfoil platform (<NUM>); and
a hollow airfoil section (<NUM>) joining the first airfoil platform (<NUM>) and the second airfoil platform (<NUM>), the hollow airfoil section (<NUM>) including a collar projection (<NUM>) extending past the first airfoil platform (<NUM>), the collar projection (<NUM>) including at least one radial tab (<NUM>) projecting therefrom;
a spar piece (<NUM>) defines a spar platform (<NUM>) and a spar extends from the spar platform (<NUM>) into the hollow airfoil section (<NUM>), the spar piece (<NUM>) including a radial pocket (<NUM>) defined by first and second opposed faces, the radial pocket (<NUM>) configured to receive the collar projection (<NUM>); and
characterized in that:
the hollow airfoil (<NUM>) includes first and second cavities (<NUM>) separated by a divider (<NUM>), and wherein the at least one radial tab (<NUM>) is at a circumferential location of collar projection (<NUM>) that corresponds to a location of the divider (<NUM>).