Patent Description:
Ceramics such as ceramic matrix composite ("CMC") materials are also being considered for use in various components of gas turbine engines. Among other attractive properties, CMCs have high temperature resistance and oxidation resistance. Despite these attributes, however, there are unique challenges to implementing CMCs in airfoils.

<CIT> discloses a prior art airfoil vane assembly as set forth in the preamble of claim <NUM>.

<CIT> discloses a prior art turbine vane assembly with cooling features.

From one aspect, there is provided an airfoil vane assembly as recited in claim <NUM>.

There is also provided a gas turbine engine as recited in claim <NUM>.

There is also provided a method of making a spar piece for an airfoil vane assembly as recited in claim <NUM>.

In a further example of the foregoing, the method includes masking the sealing surface prior to the applying such that the thermal barrier coating is not deposited on the sealing surface.

The engine parameters described above and those in this paragraph are measured at this condition unless otherwise specified. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about <NUM>, or more narrowly greater than or equal to <NUM>. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about <NUM> ft / second (<NUM> meters/second), and can be greater than or equal to <NUM> ft / second (<NUM> meters/second).

Ceramic materials are of interest for use in various components of gas turbine engines <NUM> due to their high heat tolerance and good oxidation resistance. The ceramic may be a monolithic ceramic or 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 a desired geometry. 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 or unidirectional, for example.

One example ceramic component for a gas turbine engine <NUM> is a vane. <FIG> illustrates a perspective view of 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>. <FIG> illustrates a sectioned view of the vane <NUM> along the section line shown in <FIG>. A plurality of vanes <NUM> are situated in a circumferential row about the engine central axis A (<FIG>). The vane <NUM> is comprised of a vane piece <NUM> and a spar piece <NUM>. The vane 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 platforms <NUM>/<NUM> span between a leading edge LE and a trailing edge TE. The platforms <NUM>/<NUM> The airfoil section <NUM> includes at least one internal passage <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.

As noted above, the vane piece <NUM> may be a CMC material including fiber plies that are arranged in a stacked configuration and formed to the desired geometry of the vane piece <NUM>. 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 vane piece <NUM> may be continuous in that the fiber plies are uninterrupted through the first platform <NUM>, the airfoil section <NUM>, and the second platform <NUM>. In alternate examples, the vane 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 vane 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. In some examples, the spar platform <NUM> is a first (radially outer) platform, and the spar piece <NUM> further includes a second (radially inner) platform <NUM>. The spar piece <NUM> provides mechanical support to the vane piece <NUM>. In some examples, the spar piece <NUM> may also act as a baffle to promote the flow of cooling air through the internal passage <NUM> of the vane piece <NUM>.

With continued reference to <FIG>, one or more seals <NUM> are provided to seal off the internal passage <NUM>. As noted above, the internal passage <NUM> receives cooling air flow F. The seals <NUM> maintain pressure of cooling air flow F within the internal passage <NUM> in order to maximize cooling effects of the cooling air flow F on the vane piece <NUM> and spar piece <NUM>. The seals <NUM> could be any type of seal that is known in the art. In the example of <FIG>, there are four seals <NUM>, one seal at each of the trailing edge TE and the leading edge LE of the interface between outer platform <NUM>/spar platform <NUM> and inner platform <NUM>/spar platform <NUM>. More or less seals could be used in other examples. The seals <NUM> are arranged at sealing surfaces <NUM> of the platforms <NUM>/<NUM> and span to the platforms <NUM>/<NUM>.

As noted above, ceramic-based components such as the vane piece <NUM> exhibit high temperature and oxidation resistance which can protect support structures such as the spar piece <NUM> from heat and oxidation. Additionally, because ceramic-based components have lower cooling requirements as compared to metallic components, a lower outflow margin (e.g., ratio of air pressure inside the component to air pressure outside the component) for cooling air can be used with ceramic components.

Moreover, the support structures such as the spar piece <NUM> are often metallic and therefore do not have the same inherent temperature and oxidation resistant properties. In the event of damage to ceramic-based components, the support structures could experience excessive heating because of the loss of heat protection from the ceramic-based component combined with the lower outflow margin employed for some ceramic-based components, as discussed above. In some examples, the support structures can carry cooling air to other parts of the engine <NUM>, so damage to the ceramic-based components, and subsequent excessive heating of the support structures, could interfere with the flow of cooling air throughout the engine <NUM>. In some extreme examples, the support structures could themselves become susceptible to damage from heat or oxidative effects.

Accordingly, the support structure includes a thermal barrier coating ("TBC"). The TBC improves the capability of the support structure to withstand high temperature environments as is known in the art, irrespective of the state or presence of the ceramic-based component. Moreover, even with the presence of an undamaged ceramic-based component, the TBC allows for use of even lower outflow margins because the additional heat protection provided by the TBC decreases the cooling requirements for the support structure. With a lower outflow margin, there is less risk of leakage across seals such as seals <NUM> since the air pressure within the ceramic component/support structure is decreased, leading to overall improved cooling efficiency within the engine <NUM>.

With continued reference to <FIG>, the spar piece <NUM> includes a TBC <NUM>. Various TBCs are known in the art, and any TBC <NUM> could be employed. In general, TBCs include a bond coat and at least one ceramic layer. Moreover, various methods of TBC application to a metallic component are known in the art, and any method of application could be used. In some examples, the TBC <NUM> has abrasion-resistance properties in addition to heat-resistance properties. The TBC <NUM> is disposed on at least the spar <NUM>, but it can also extend along at least a portion of the spar platforms <NUM>/<NUM>. However, the sealing surfaces <NUM> are free from TBC <NUM>. Because the sealing surfaces <NUM> are free from TBC <NUM>, the TBC <NUM> does not interfere with the sealing effectiveness of the seals <NUM>.

In some examples, a clearance L is provided between the end of the TBC <NUM> and the seal <NUM>. The clearance L can be up to about <NUM> mil (<NUM>) long. For example, prior to application of the TBC <NUM> to the spar piece <NUM>, the sealing surface <NUM> and desired clearance L could be masked off to prevent deposition of TBC <NUM> in the sealing surface <NUM> and/or clearance L. Masking is a well-known procedure that is compatible with various coating deposition methods such as air plasma spraying or others.

As shown in <FIG>, the TBC <NUM> includes a taper <NUM> at the end of the TBC <NUM> as the TBC <NUM> approaches the seal <NUM>. In this example, the length T of the taper can be up to about <NUM> mils (<NUM>). The clearance L can optionally be employed together with the taper <NUM>.

In some examples, the TBC <NUM> has a thickness that is less than about <NUM> mils (<NUM>). In a further example, the TBC <NUM> has a thickness that is less than about <NUM> mils (<NUM>). In a further example, the TBC <NUM> has a thickness that is between about <NUM> mils (<NUM>) and about <NUM> mils (<NUM>). In general, the TBC <NUM> does not substantially change the size of the spar piece <NUM> and does not substantially decrease the size of the clearance between the spar <NUM> and the airfoil section <NUM>.

<FIG> illustrate example blade outer air seals (BOAS). Like vanes <NUM>, BOAS can also employ ceramic-based components with metallic support structures. Accordingly, BOAS can also benefit from the TBC as discussed above. In the example of <FIG>, a ceramic-based BOAS <NUM> is supported on a metallic carrier <NUM>. The ceramic-based BOAS <NUM> could be a monolithic ceramic material or a CMC material, as discussed above. The carrier <NUM> is in turn connected to a casing structure <NUM> of the engine <NUM>. The carrier <NUM> includes a TBC <NUM> on its radially inner side (e.g., the side adjacent the BOAS <NUM>). In this example, a clearance L can be provided between the TBC <NUM> and BOAS <NUM>. Alternatively or additionally, the TBC <NUM> can have abrasion-resistant properties, as discussed above.

In the example of <FIG>, a ceramic-based BOAS <NUM> is supported on supports <NUM> of a casing structure <NUM> of the engine <NUM> by a connector <NUM>. The ceramic-based BOAS <NUM> could be a monolithic ceramic material or a CMC material, as discussed above. In this example, the casing structure <NUM> includes TBC <NUM>, however, interfaces <NUM> of the supports <NUM>/connector <NUM> are free of TBC <NUM> much like sealing surfaces <NUM> in the example of <FIG> are free from TBC <NUM>. The same clearance L discussed above can be employed in this example between end of the TBC <NUM> and the interfaces <NUM>. Alternatively or additionally, a taper <NUM> as shown above in <FIG>.

In the examples of Figures 3A-C, the TBC <NUM> can have a thickness that is less than about <NUM> mils (<NUM>). In a further example, the TBC <NUM> has a thickness that is less than about <NUM> mils (<NUM>). In a further example, the TBC <NUM> has a thickness that is between about <NUM> mils (<NUM>) and about <NUM> mils (<NUM>).

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:
An airfoil vane assembly (<NUM>), comprising:
a vane piece (<NUM>) having a first vane platform (<NUM>), a second vane platform (<NUM>), and a hollow airfoil section (<NUM>) joining the first vane platform (<NUM>) and the second vane platform (<NUM>);
a spar piece (<NUM>) having a spar platform (<NUM>) and a spar (<NUM>) extending from the spar platform (<NUM>) into the hollow airfoil section (<NUM>);
at least one seal (<NUM>) arranged at a sealing surface (<NUM>) of the spar platform (<NUM>) and sealing between the spar platform (<NUM>) and the first vane platform (<NUM>); and
a thermal barrier coating (<NUM>) disposed on the spar piece (<NUM>), wherein the sealing surface is free from the thermal barrier coating (<NUM>).,
characterised in that
the thermal barrier coating (<NUM>) extends along a portion of the spar platform (<NUM>) and includes a taper (<NUM>) at an end of the thermal barrier coating (<NUM>), as the thermal barrier coating (<NUM>) approaches the seal (<NUM>).