Patent Description:
Gas turbine engines are known, and typically include a fan delivering air into a bypass duct as propulsion air. The air is also delivered into a compressor where it is compressed. The compressed air is delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving them to rotate. The turbine rotors in turn drive compressor and fan rotors.

As known, the turbine sees very high temperatures from the products of combustion. As such, it has been proposed to use ceramic matrix composites ("CMC") for various components as they are better able to withstand high temperature.

It is also desirable to ensure that the great bulk of the products of combustion pass over blades on the turbine rotors to drive them to rotate. This improves the efficiency. Thus, it is known to provide seals to limit leakage around the turbine blades.

One component which is often provided to limit leakage is a blade outer air seal ("BOAS"). The blade outer air seal is positioned radially outwardly of a tip of a turbine blade. Typically, a blade outer air seal has a leading edge mount hook and a trailing edge mount hook.

A chamber is defined between the two mount hooks. Cooling air is directed into this chamber, and typically from a compressor associated with the gas turbine engine. This cooling air is at a relatively higher pressure, and at a high pressure than the products of combustion. As such, there is a tendency for the high pressure cooling air to leak around the BOAS and into the products of combustion. This creates inefficiencies and it is undesirable.

Thus, it is known to provide brush seals between a static housing and a radially outer or back side of the blade outer air seal to limit this leakage.

In general, the brush seal has extended radially inwardly with only a tangential angle to prevent buckling, and axially perpendicular to a rotational axis of the engine to contact the back of the blade outer air seal.

Static turbine vanes are also positioned intermediate turbine blade rows. It is known to provide brush seals on such vanes.

In at least one known prior art vane, a back surface of the vane was formed at a non-parallel angle relative to the rotational axis of the associated engine. In this vane, a brush seal extended radially inwardly at an angle that was not perpendicular to the rotational axis. In the location where the brush seal contacted the back surface of the vane, the brush seal was essentially at a perpendicular angle relative to the back surface.

In another prior art arrangement the brush seal extended radially inwardly with an axial component in a downstream direction.

<CIT> discloses a seal assembly with an impingement seal plate.

According to a first aspect of the invention, there is provided a gas turbine engine as claimed in claim <NUM>.

In an embodiment, the angle is between <NUM> and <NUM> degrees.

In another embodiment according to any of the previous embodiments, the sealing surface is defined downstream of the trailing edge attachment feature, and upstream of the trailing edge with the axial component being towards the trailing edge attachment feature such that the angle is on a downstream side of the bristles.

In another embodiment according to any of the previous embodiments, there is a second of the brush seals, with a second of the sealing surfaces defined upstream of the leading edge attachment feature, and downstream of the leading edge, with the axial component being toward the leading edge attachment feature such that the angle is on an upstream side of the bristles.

In another embodiment according to any of the previous embodiments, the sealing surface is defined upstream of the leading edge attachment feature, and downstream of the leading edge, with the axial component being toward the leading edge attachment feature such that the angle is on an upstream side of the bristles.

In another embodiment according to any of the previous embodiments, a circumferential direction is defined for <NUM> degrees about the rotational axis. The bristles are wire bristles which are circumferential positioned at a non-parallel angle relative to a radial direction extending outwardly from the axis of rotation.

In another embodiment according to any of the previous embodiments, the sealing surface is generally parallel, or parallel, to the rotational axis.

In another embodiment according to any of the previous embodiments, the sealing surface extends at an angle which is non-parallel to the rotational axis.

In another embodiment according to any of the previous embodiments, the bristles are mounted to extend generally radially inwardly, or radially inward, relative to the rotational axis to contact the sealing surface and define the angle.

In another embodiment according to any of the previous embodiments, the blade outer air seal is formed of ceramic matrix composites.

According to another aspect of the invention, there is provided a gas turbine engine component as claimed in claim <NUM>.

In another embodiment according to any of the previous embodiments, the angle is between <NUM> and <NUM> degrees.

In another embodiment according to any of the previous embodiments, the blade outer air seal has a leading edge mount hook adjacent an upstream leading edge of the blade outer air seal and a trailing edge mount hook adjacent a downstream trailing edge of the blade outer air seal.

In another embodiment according to any of the previous embodiments, the sealing surface is defined downstream of the trailing edge mount hook, and upstream of the trailing edge.

In another embodiment according to any of the previous embodiments, the sealing surface extends at an angle which is non-parallel to the rotational axis. The bristles are mounted to extend generally radially inwardly, or radially inward, to contact the sealing surface relative to the rotational axis.

The fan section <NUM> may include a single-stage fan <NUM> having a plurality of fan blades <NUM>. The fan blades <NUM> may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan <NUM> drives air along a bypass flow path B in a bypass duct <NUM> defined within a housing <NUM> such as a fan case or nacelle, and also drives air along a core flow path C for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. A splitter <NUM> aft of the fan <NUM> divides the air between the bypass flow path B and the core flow path C. The housing <NUM> may surround the fan <NUM> to establish an outer diameter of the bypass duct <NUM>. The splitter <NUM> may establish an inner diameter of the bypass duct <NUM>. Although depicted as a two-spool turbofan gas turbine engine in the disclosed nonlimiting 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 three-spool architectures. The engine <NUM> may incorporate a variable area nozzle for varying an exit area of the bypass flow path B and/or a thrust reverser for generating reverse thrust.

The inner shaft <NUM> is connected to the fan <NUM> through a speed change mechanism, which in the exemplary gas turbine engine <NUM> is illustrated as a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The inner shaft <NUM> may interconnect the low pressure compressor <NUM> and low pressure turbine <NUM> such that the low pressure compressor <NUM> and low pressure turbine <NUM> are rotatable at a common speed and in a common direction. In other embodiments, the low pressure turbine <NUM> drives both the fan <NUM> and low pressure compressor <NUM> through the geared architecture <NUM> such that the fan <NUM> and low pressure compressor <NUM> are rotatable at a common speed. Although this application discloses geared architecture <NUM>, its teaching may benefit direct drive engines having no geared architecture.

The low pressure compressor <NUM>, high pressure compressor <NUM>, high pressure turbine <NUM> and low pressure turbine <NUM> each include one or more stages having a row of rotatable airfoils. Each stage may include a row of vanes adjacent the rotatable airfoils. The rotatable airfoils are schematically indicated at <NUM>, and the vanes are schematically indicated at <NUM>.

The engine <NUM> may be a high-bypass geared aircraft engine. The bypass ratio can be greater than or equal to <NUM> and less than or equal to about <NUM>, or more narrowly can be less than or equal to <NUM>. The geared architecture <NUM> may be an epicyclic gear train, such as a planetary gear system or a star gear system. The epicyclic gear train may include a sun gear, a ring gear, a plurality of intermediate gears meshing with the sun gear and ring gear, and a carrier that supports the intermediate gears. The sun gear may provide an input to the gear train. The ring gear (e.g., star gear system) or carrier (e.g., planetary gear system) may provide an output of the gear train to drive the fan <NUM>. A gear reduction ratio may be greater than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, and in some embodiments the gear reduction ratio is greater than or equal to <NUM>. The fan diameter is significantly larger than that of the low pressure compressor <NUM>. The low pressure turbine <NUM> can have a pressure ratio that is greater than or equal to <NUM> and in some embodiments is greater than or equal to <NUM>. All of these parameters are measured at the cruise condition described below.

The engine parameters described above, and those in the next paragraph are measured at this condition unless otherwise specified.

"Fan pressure ratio" is the pressure ratio across the fan blade <NUM> alone, without a Fan Exit Guide Vane ("FEGV") system. A distance is established in a radial direction between the inner and outer diameters of the bypass duct <NUM> at an axial position corresponding to a leading edge of the splitter <NUM> relative to the engine central longitudinal axis A. The fan pressure ratio is a spanwise average of the pressure ratios measured across the fan blade <NUM> alone over radial positions corresponding to the distance. The fan pressure ratio can be less than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, such as between <NUM> and <NUM>. "Corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]<NUM>. The corrected fan tip speed can be less than or equal to <NUM> ft / second (<NUM> meters/second), and can be greater than or equal to <NUM> ft / second (<NUM> meters/second).

<FIG> shows a portion of a turbine section <NUM>. A blade outer air seal <NUM> has a radially inward facing surface <NUM> closely spaced from a radially outer tip <NUM> of a rotating turbine blade <NUM>.

The blade outer air seal <NUM> has a leading edge mount hook <NUM> associated with a leading edge <NUM>. The blade outer air seal <NUM> also has a trailing edge mount hook <NUM> associated with a trailing edge <NUM>.

As is known, products of combustion pass across the turbine blade <NUM> from its leading edge <NUM>, or an upstream location toward the trailing edge <NUM> or downstream. The blade outer air seal <NUM> serves to limit leakage of the products of combustion around the turbine blade <NUM>.

There is a face <NUM> on a radial side of the blade outer air seal <NUM> that is remote from the face <NUM>. The face <NUM> has a surface <NUM> between the mount hooks <NUM> and <NUM>. There is also a surface <NUM> which is between the trailing edge mount hook <NUM> and the trailing edge <NUM>. Similarly, there is a surface <NUM> between the leading edge mount hook <NUM> and the leading edge <NUM>. The blade outer air seal <NUM> is mounted on mount structure <NUM>, shown schematically.

As shown, a source <NUM>, which may be tapped from a compressor associated with the gas turbine engine such as shown in <FIG>, communicates high pressure cooling air into a chamber <NUM> defined between the leading edge mount hook <NUM> and the trailing edge mount hook <NUM>. Typically, the cooling air flows into a cooling circuit in a body of the blade outer air seal. This high pressure cooling air is at a higher pressure than the products of combustion in a chamber <NUM>. As such, there is a tendency for the high pressure cooling air to leak around the hooks <NUM> and <NUM> and mix into the products of combustion. This would be inefficient.

While mount hooks are disclosed, blade outer air seals having other types of attachment features would benefit from the teaching of this disclosure. As other examples the attachment features could be flanges of dovetail joints.

As such, it is known to mount brush seals to resist that leakage flow. As shown in <FIG>, there is a first brush seal <NUM> located to seal on the surface <NUM>. There is a second brush seal <NUM> to seal on the surface <NUM>. As shown, the seals <NUM>//<NUM> have a housing <NUM> and <NUM> which is formed of two pieces to mount a plurality of brush bristles <NUM>, or bristle pack. Mount structure <NUM> is shown schematically mounting the brush seals <NUM> and <NUM>.

While two brush seals are shown at two distinct locations, embodiments having only a single brush seal at either of these locations may benefit from this disclosure. Further, additional brush seals may be utilized if indicated.

As shown in <FIG>, the bristles <NUM> extend to a tip <NUM> which is in contact with the surfaces <NUM> and <NUM>. As illustrated, and as will be explained below, the bristles <NUM> are not perpendicular to the surface <NUM>, but rather extend along a direction that includes a component extending toward the leading edge <NUM> and having a component extending radially inwardly when extending from the mount toward the blade outer air seal surface <NUM> that contacts bristles <NUM>. Due to this, a pressure differential between a pressure P<NUM> found in the chamber <NUM>, compared to the pressure P<NUM> found in the chamber <NUM>, forces F the bristles <NUM> against the surface <NUM>. In prior art blade seals wherein the bristles extend generally perpendicular to the sealing surface, the bristles might be forced to bend inwardly, and toward the trailing edge <NUM>, allowing leakage.

As shown, the leading edge brush seal <NUM> is at an angle which is generally opposed to the angle of the trailing edge brush seal <NUM>. Still, there is a high pressure P1 from the chamber <NUM> that will force the bristles <NUM> against the surface <NUM> relative to pressure P2 in chamber <NUM>.

It could be said that the bristles extend radially inwardly from a mount location and with an axial component from the surfaces <NUM> and <NUM>, and in a direction towards the mount hooks to define an angle on a side of the brush seals <NUM> and <NUM> opposite the mount hooks <NUM>/<NUM> wherein the angle is less than <NUM>°.

<FIG> is a view along line <NUM>-<NUM>, and shows the bristles <NUM> are angled with a circumferential component as they extend around the circumference of the blade outer air seal <NUM>. Further, the blade outer air seal <NUM> may be formed by a plurality of distinct segments, as is illustrated schematically by the dashed lines <NUM>. It is preferable that the brush seal extends across the entirety of <NUM> degrees relative to a rotational axis X of the gas turbine engine.

<FIG> shows an embodiment <NUM> wherein a blade outer air seal <NUM> has its bristles <NUM> extending generally perpendicularly relative to a rotational axis X. However, the blade outer air seal <NUM> has its sealing surface <NUM> formed at a non-parallel angle relative to the rotational axis X. Here again, when the pressure differential across the blade outer air seals acts against the bristles, they will be forced into contact with the surface <NUM>. The illustrated embodiment is downstream of the trailing edge mount hook, such that there is an included angle between bristles <NUM> and surface <NUM> that is less than <NUM>°.

<FIG> shows the angle A between the bristles <NUM> and the surface <NUM>. This would be true of all of the brush seal locations <NUM>, <NUM> and the relationship as shown in the embodiment <NUM> of <FIG>. In embodiments, the angle A is preferably between <NUM>° and <NUM>°. In further embodiments, the angle A is between <NUM>° and <NUM>°.

<FIG> shows an alternative arrangement <NUM>, outside the wording of the claims, wherein the component contacting the brush seal <NUM> is a stator vane <NUM> which is positioned intermediate rotating turbine blade rows <NUM> that is less than <NUM>°. The bristles in the brush seal <NUM> are again at an angle relative to a sealing surface <NUM> on the stator vane <NUM>.

In this arrangement, a chamber <NUM> contains gasses at a higher pressure than a chamber <NUM>.

Generically, it could be said that the embodiments or arrangements of <FIG> and <FIG> all have a brush seal with bristles extending radially inwardly to seal between a first higher pressure chamber and a second lower pressure chamber, and the bristles defining an angle with a sealing surface having an axial component extending in a direction toward the higher pressure chamber.

This particular disclosure is especially powerful when utilized with components formed of ceramic matrix composites. In manufacturing components from ceramic matrix composites it is sometimes difficult to form a very smooth sealing surface. The additional sealing contact supplied by the pressure differential will thus allow the bristles in the several seals to better provide a seal on the sealing surface.

In embodiments, the ceramic matrix components could be formed of CMC material or a monolithic ceramic. A CMC material is comprised of one or more ceramic fiber plies in a ceramic matrix. Example ceramic matrices are silicon-containing ceramic, such as but not limited to, a silicon carbide (SiC) matrix or a silicon nitride (Si3N4) matrix. Example ceramic reinforcement of the CMC are silicon-containing ceramic fibers, such as but not limited to, silicon carbide (SiC) fiber or silicon nitride (Si3N4) fibers. The CMC may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fiber plies are disposed within a SiC matrix. A fiber ply has a fiber architecture, which refers to an ordered arrangement of the fiber tows relative to one another, such as a 2D woven ply or a 3D structure. A monolithic ceramic does not contain fibers or reinforcement and is formed of a single material. Example monolithic ceramics include silicon-containing ceramics, such as silicon carbide (SiC) or silicon nitride (Si3N4).

Claim 1:
A gas turbine engine (<NUM>) comprising:
a compressor section (<NUM>) for receiving air and delivering it to a combustor (<NUM>);
a turbine section (<NUM>, <NUM>) downstream of said combustor (<NUM>), an upstream location leading from the compressor section (<NUM>) in a downstream direction toward the turbine section (<NUM>, <NUM>), the turbine section (<NUM>, <NUM>) including at least one rotating blade row (<NUM>) for rotation about a rotational axis (X), and a blade outer air seal (<NUM>) positioned radially outward of blades in the rotating blade row (<NUM>) relative to the rotational axis (X), the blade outer air seal (<NUM>) having a radially inward facing web (<NUM>), spaced from a radially outer tip (<NUM>) of the blades, and a radially outer face (<NUM>) relative to said radially inward facing web (<NUM>), wherein said blade outer air seal (<NUM>) has a leading edge attachment feature (<NUM>) adjacent an upstream leading edge (<NUM>) of the blade outer air seal (<NUM>) and a trailing edge attachment feature (<NUM>) adjacent a downstream trailing edge (<NUM>) of the blade outer air seal (<NUM>); and
at least one brush seal (<NUM>; <NUM>) having bristles (<NUM>) extending radially inward from a mount location to contact a sealing surface (<NUM>; <NUM>) on the radially outer face (<NUM>), wherein a chamber (<NUM>) is defined between said leading edge attachment feature (<NUM>) and said trailing edge attachment feature (<NUM>), and adapted to be connected to a source (<NUM>) of pressurized cooling air, with an angle (A) defined between said bristles (<NUM>) and said sealing surface (<NUM>; <NUM>) having a radially inward extending component, and with an axial component in a direction toward said leading edge attachment feature (<NUM>) and said trailing edge attachment feature (<NUM>), and the angle (A) is less than <NUM>° on a side of the brush seal (<NUM>; <NUM>) opposite the leading edge and trailing edge attachment features (<NUM>, <NUM>).