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-pressure and temperature gas flow. The high-pressure and temperature gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section may include low and high pressure compressors, and the turbine section may also include 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 capability and lifetime. Ceramic matrix composite ("CMC") materials are also being considered for airfoils. Among other attractive properties, CMCs have high temperature resistance. Despite this attribute, however, there are unique challenges to implementing CMCs in airfoils.

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

<CIT> discloses a prior art cooling arrangement for engine components.

<CIT> discloses prior art manufacture of full ring strut vane pack.

<CIT> discloses a prior art stator damper and corresponding method of fabrication.

<CIT> discloses a prior art ceramic guide vane for a gas turbine engine.

<CIT> discloses a prior art stator blade ring and axial flow compressor using the same.

<CIT> discloses a prior art elastic fluid turbine bucket wheel.

From one aspect, there is provided a seal assembly as recited in claim <NUM>.

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

The various features and advantages of the present invention will become apparent to those skilled in the art from the following detailed description.

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).

<FIG> illustrates an example implementation of a seal assembly <NUM> in the turbine section <NUM> of the engine <NUM> (see also <FIG>). The seal assembly <NUM> is employed with a plurality of gas turbine engine components <NUM> (one shown in <FIG>) that are arranged circumferentially around the central engine axis A. The gas turbine engine components <NUM> are vane assemblies.

The vane assemblies <NUM> each include an airfoil fairing <NUM> that is comprised of an airfoil section <NUM> and first and second platforms <NUM>/<NUM> between which the airfoil section <NUM> extends. The platform <NUM> includes a base wall 66a, forward and aft flanges 66b/66c that protrude radially from the base wall 66a, and mate faces 66d. The mate faces 66d are the circumferential sides of the platform <NUM> and may include a slot <NUM> for receiving a feather seal (not shown). Brush seals <NUM> are provided at the forward and trailing ends of the platform <NUM> to facilitate isolating a plenum region bounded by the base wall 66a and the flanges 66b/66c. In this example, the platform <NUM> also includes a collar 66e that protrudes radially from the base wall 66a. Although not shown, the platform <NUM> may also have one or more flanges, which may serve to transfer aerodynamic loads from the airfoil fairing <NUM>. In addition, platform <NUM> may have protruding flanges instead or, or in addition to, the collar. In the example shown in <FIG> the first platform <NUM> is an outer platform and the second platform <NUM> is an inner platform. The terms such as "inner" and "outer" used herein refer to location with respect to the central engine axis A, i.e., radially inner or radially outer. Moreover, the terminology "first" and "second" 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 that a first component or feature could alternatively be termed as the second component or feature, and vice versa.

The airfoil section <NUM> generally extends in a radial direction relative to the central engine axis A and defines leading and trailing edges 64a/64b, a suction side 64c, and a pressure side 64d. The airfoil section <NUM> is hollow and circumscribes an interior through-cavity <NUM>. The airfoil section <NUM> may have a single through-cavity <NUM>, or the cavity <NUM> may be sub-divided by one or more ribs. In the example shown, the airfoil fairing includes one rib 70a that sub-divides the cavity <NUM> into forward and aft sub-cavities. The aforementioned collar 66e is an extension of a tube that forms the rib 70a and the forward sub-cavity in the airfoil section <NUM>.

The airfoil fairing <NUM> is a continuous, one-piece body. As an example, the airfoil fairing <NUM> is formed of a ceramic material or a metal matrix composite (MMC). For instance, the material is a ceramic matrix composite, or a metal matrix composite (MMC). In one example, the ceramic matrix composite (CMC) is formed of ceramic fiber tows that are disposed in a ceramic matrix. The ceramic matrix composite may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fiber tows are disposed within a SiC matrix. Example metal matrix composites include, but are not limited to, boron carbide fiber tows and/or alumina fiber tows disposed in a metal matrix, such as aluminum. The fiber tows are arranged in a fiber architecture, which refers to an ordered arrangement of the tows relative to one another, such as a 2D woven ply (e.g. a braid) or a 3D structure.

The vane assembly <NUM> further includes a spar <NUM> that extends through the through-cavity <NUM> (e.g., the forward sub-cavity) and mechanically supports the airfoil fairing <NUM>. The spar <NUM> may be formed of a relatively high temperature resistance, high strength material, such as a nickel or cobalt based superalloy (e.g. a single crystal nickel alloy). The spar <NUM> includes a spar flange 72a and a spar leg 72b that extends from the spar flange 72a, through the collar 66e of the platform <NUM>, and into the through-cavity <NUM> (e.g., the forward sub-cavity). The spar leg 72b defines an interior through-passage 72c, and the spar flange 72a is supported at an outer end to support hardware S.

The spar leg 72b has an inner end portion that has an attachment <NUM>, such as but not limited to, a pin. The inner end of the spar leg 72b extends past the platform <NUM> of the airfoil fairing <NUM> so as to protrude from the fairing <NUM>. The inner end is attached (e.g., by the pin) to additional support hardware S adj acent the platform <NUM> of the airfoil fairing <NUM>. The airfoil fairing <NUM> is thus trapped between the inner support hardware S and the spar flange 72a and outer support hardware S.

Airflow, such as bleed air from the compressor section <NUM>, is provided in and around the vane assembly to meet various objectives. For instance, in zone Z1 air is used to purge the region around the forward flange 66b and brush seal <NUM>. In zone Z2 air is used to cool the region of the leading edge 64a of the airfoil section <NUM>; and in zone Z3 air may be used to purge the region around aft flange 66c. The pressures across the zones Z1/Z2/Z3 are different and to the extent that air is permitted to leak from the zones, the above objectives may be frustrated. Moreover, since zone Z3 is at the lowest pressure, air from the zones Z1 and Z2 tends to leak to zone Z3. In this regard, the seal assembly <NUM> facilitates isolation of the zones Z1/Z2/Z3.

The seal assembly <NUM> is implemented at the outer platforms <NUM> of the airfoil fairings <NUM> and includes a seal arc segment <NUM>. <FIG> illustrates an isolated view of the seal arc segment <NUM>. As the phrase "seal arc" indicates, the seal arc segment <NUM> is a sector of a circular ring. In this regard, the seal arc segment <NUM> spans across at least three or more vane assemblies <NUM>. For instance, multiple seal arc segments <NUM> would make a full ring, such as bi-segment halves, tri-segment thirds, or quad-segment quarters in a ring.

Each seal arc segment <NUM> includes seal portions 76a that are circumferentially spaced-apart and a connector portions 76b that join the seal portions 76a such that there are alternating seal portions 76a and connector portions 76b. In this example, each of the connector portions 76b includes an opening 76c to receive the collar 66e there through. The opening 76c is completely bound within the connector portion 76b, and the shape of the opening 76c is a complement of the shape of the collar 66e such that the collar 66e fits closely through the opening 76c.

The seal arc segment <NUM> may be formed of a metal alloy, such as a nickel- or cobalt-based alloy, or a ceramic matrix composite. In one example, the airfoil fairing <NUM> and the seal arc segment <NUM> are both formed of CMCs. In a further example, the CMCs of the airfoil fairing <NUM> and the seal arc segment <NUM> are of equivalent compositions. For instance, the equivalent compositions have the same kind of fibers (e.g., silicon carbide) and the same kind of matrix (e.g., silicon carbide) however the layup architecture (weave, braid, uni-tape, etc.) may be different.

As shown in <FIG>, the vane assemblies <NUM> are circumferentially successively arranged such that the (first) mate face 66d of the first one of the vane assemblies (the left-most) is adjacent the (second) mate face 66d of the second one of the vane assemblies (the middle) to define a (first) mate face gap <NUM> there between. The (first) mate face 66d of the second (middle) one of the vane assemblies <NUM> is adjacent the (second) mate face 66d of the third one of the vane assemblies (rightmost) to define a (second) mate face gap <NUM> there between.

The seal arc segment <NUM> is received onto the platforms <NUM> of the three circumferentially successive vane assemblies <NUM>. The collars 66e fit into the openings 76c in the connector portions 76b of the seal arc segments <NUM>, which may also serve as pilots for proper location of the seal arc segment <NUM>. The seal arc segment <NUM> is situated on the base walls 66a of the platforms <NUM> (see also <FIG>) axially between the flanges 66b/66c such that the connector portions 76b span circumferentially across the base walls 66a. The seal portions 76a contact the base walls 66a of the adjacent platforms <NUM> and bridge across the mate face gaps <NUM> to seal the mate face gaps <NUM>. The brush seals <NUM> (<FIG>) contact the outer surface of the seal arc segment <NUM> to seal therewith.

Although feather seals may be used in the slots <NUM> for sealing the mate faces 66d, the seal portions 76a that bridge over the mate face gaps <NUM> provide additional sealing against infiltration of hot gas from the core gas path and/or the escape of cooling air from zone Z2. This additional sealing, together with the sealing provided by the brush seals <NUM>, thereby facilitates the isolation of zone Z2 from zones Z1 and Z3. The seal arc segment <NUM> may also serve as a thermal shield to protect the brush seals <NUM> from high temperatures that may be experienced in the platform <NUM>. In this regard, forming the seal arc segment <NUM> from a relatively low thermal conductivity material (in comparison to metals), such as the CMC discussed above, facilitates enhancement of thermal shielding. Alternatively, a metal arc seal segment may be coated with a thermal barrier coating such as yttria stabilized zirconia.

As also shown in <FIG>, the mate face gaps <NUM> each define a mate face gap length G, which is equal to the axial length of the platforms <NUM>. As the seal arc segments <NUM> are located between the flanges 66b/66c, the seal portions 76a extend only along a majority of the mate face gap length G rather than the full extent of the mate face gap length G. For instance, as shown in <FIG>, the seal portions 76a extend over the middle section of the mate face gap length G and opposed end sections of the mate face gap length G extend forward of and aft of the seal portions 76a. Thus, the seal portions 76a do not seal the entire mate face gap <NUM>, and the end sections of the mate face gap length G rely on only the sealing provided by the feather seals.

Optionally, as shown in <FIG>, the seal portions 76a may include one or more impingement holes 76d. Cooling air that is provided into zone Z2 (see <FIG>) is discharged through the impingement holes 76d into the mate face gaps <NUM>. The size and number of impingement holes 76d may be selected to meter the amount of cooling air provided, to cool the feather seal and/or cool the mate faces 66d. The impingement holes 76d may also be placed in other regions of the seal portion 76b to cool specific areas of the platform <NUM>.

Optionally, the perimeter surface regions of the fairing platforms <NUM> that are in contact with the seal portions 76a may include a coating, for thermal considerations, to smooth the surface for better sealing contact, and/or reduce wear. For example, the coating is selected of a composition that is thermally insulating in comparison to the CMC (if used) of the airfoil fairings <NUM>, to thermally insulate the seal arc segment <NUM>. The coating may be composed of elemental silicon, silicate, silica, hafnia, zirconia, or combinations thereof. As the coating may be only on the perimeter portions of the platforms, including along the mate face edges 66d, there may be a cavity <NUM> defined between the seal arc segment <NUM>, the radial surface of the platform <NUM>, and the coating. One or more impingement holes 76d may be provided in the connector portions 76b that open to the cavity <NUM> to provide cooling to the surface of the platform <NUM> and thereby facilitate control over the temperature of regions in contact with the brush seals <NUM>.

Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this invention.

In other words, a system designed according to an embodiment of this invention will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures.

Claim 1:
A seal assembly (<NUM>) comprising:
first, second, and third gas turbine engine components (<NUM>) that are successively arranged around an axis (A), each of the first, second, and third gas turbine engine components (<NUM>) being a vane assembly having a first and a second platform (<NUM>, <NUM>) between which an airfoil section (<NUM>) extends, the first platform (<NUM>) having first and second mate faces (66d) such that the first mate face (66d) of the first gas turbine engine component (<NUM>) is adjacent to the second mate face (66d) of the second gas turbine engine component (<NUM>) to define a first mate face gap (<NUM>) therebetween and the first mate face (66d) of the second gas turbine engine component (<NUM>) is adjacent to the second mate face (66d) of the third gas turbine engine component (<NUM>) to define a second mate face gap (<NUM>) therebetween; and
a seal arc segment (<NUM>) arranged around the axis (A), the seal arc segment (<NUM>) having first and second seal portions (76a) that are circumferentially spaced-apart and a connector portion (76b) joining the first and second seal portions (76a), the seal arc segment (<NUM>) being arranged such that the first seal portion (76a) bridges the first mate face gap (<NUM>) to seal the first mate face gap (<NUM>), the second seal portion (76a) bridges the second mate face gap (<NUM>) to seal the second mate face gap (<NUM>), and the connector portion (76b) spans circumferentially across the second gas turbine engine component (<NUM>),
characterised in that
the seal assembly (<NUM>) further comprises two brush seals (<NUM>) extending from support hardware (S) and contacting a radially outer surface of the seal arc segment (<NUM>) at a leading end of the seal arc segment (<NUM>) and at a trailing end of the seal arc segment (<NUM>) to thereby isolate a cooling air zone (Z2), wherein the cooling air zone (Z2) is configured to use air to cool the region of a leading edge (64a) of the airfoil section (<NUM>), wherein the brush seals (<NUM>) seal against the radially outer surface of the seal arc segment (<NUM>).