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.

The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction.

<CIT> discloses a prior art gas turbine transition piece seal system.

<CIT> discloses a prior art turbine system comprising a transition duct with a flexible seal.

<CIT> discloses a prior art C-shaped seal ring.

<CIT> discloses prior art scalloped cooling of gas turbine transition piece frame.

In accordance with a first aspect of the present invention, there is provided a gas turbine engine as claimed in claim <NUM>.

In an embodiment of the foregoing embodiment, the platform is radially inwards of the airfoil section.

In a further embodiment of any of the foregoing embodiments, the lip of the combustor abuts the leading edge of the platform.

In a further embodiment of any of the foregoing embodiments, the radially outer wall of the first annular slot and the radially outer wall of the second annular slot abut, and the radially inner wall of the first annular slot and the radially inner wall of the second slot are axially spaced apart such that there is a gap therebetween.

In a further embodiment of any of the foregoing embodiments, the airfoil section of each vane extends along a radial axis from the platform. Each vane has rotational play about the respective radial axis under aerodynamic loads such that each vane moves relative to the combustor between a seated state in which the forward edge abuts the lip of the combustor wall and an unseated state in which there is a divergent gap between the forward edge and the lip of the combustor. The annular feather seal is wider than the divergent gap to maintain sealing when in the vane is in the unseated state.

In an embodiment of the foregoing embodiment, the first side of the airfoil section is a suction side, the second side of the airfoil section is a pressure side, the second circumferential side of the platform is located to the second side of the airfoil section, and the divergent gap diverges toward the second circumferential side.

In a further embodiment of any of the foregoing embodiments, the seal slot is radially thicker than the annular feather seal.

"Low 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> (where °R = K x <NUM>/<NUM>).

<FIG> illustrates a view taken through selected portions of the combustor <NUM> and high pressure turbine <NUM> of the engine <NUM>. In this example, the combustor <NUM> is an annular combustor that extends around the engine central axis A, although it is contemplated that the examples herein are also applicable to can type combustors. The combustor <NUM> includes a radially inner shell or combustor wall <NUM> and a radially outer shell or combustor wall <NUM>. The walls <NUM>/<NUM> define an annular combustor chamber <NUM> there between. The combustor <NUM> includes one or more injectors <NUM> at a forward end of the combustor <NUM>, and an exit region <NUM> at the aft end of the combustor <NUM>. The combustor walls <NUM>/<NUM> each include a lip <NUM> at the axially trailing end thereof.

The high pressure turbine <NUM> includes a circumferential row of turbine vanes <NUM> adjacent the exit region <NUM>. Each vane <NUM> includes an inner or first platform <NUM>, an outer or second platform <NUM>, and an airfoil section <NUM> that spans in a radial direction between the first and second platforms <NUM>/<NUM>. A radial view of a portion of the row of turbine vanes <NUM> is also shown in <FIG>. Terms such as "radially," "axially," "circumferentially," or variations thereof are used herein to designate directionality with respect to the engine central axis A.

The airfoil section <NUM> includes an airfoil outer wall <NUM> that delimits the profile of the airfoil section <NUM>. The outer wall <NUM> defines a leading end 80a, a trailing end 80b, and first and second sides 80c/80d that join the leading and trailing ends 80a/80b. The first and second sides 80c/80d span in the radial direction between first and second ends 80e/80f that are attached, respectively, to the first and second platforms <NUM>/<NUM>. In this example, the first side 80c is a suction side and the second side 80d is a pressure side.

The first platform <NUM> (<FIG>) defines forward and trailing edges 74a/74b and first and second circumferential side edges 74c/74d that join the forward and trailing edges 74a/74b. Generally, the first and second circumferential side edges 74c/74d mate with or bear against the respective second and first circumferential side edges 74c/74d of the adjacent vanes <NUM>. Likewise, the second platform <NUM> defines forward and trailing edges and first and second circumferential side edges that join the forward and trailing edges. The first and second circumferential side edges also mate with or bear against the respective second and first circumferential side edges of the adjacent vanes <NUM>. The examples herein below may be directed to the first platform <NUM>. However, it is to be understood that the examples are also applicable to the second platform <NUM>.

<FIG> illustrates a magnified view of the exit region <NUM> of the combustor <NUM> at the forward edge 74a of the vane <NUM>, and <FIG> illustrates a sectioned view of the same area. The forward edge 74a is adjacent the lip <NUM> of the combustor wall <NUM>. In <FIG>, the vane <NUM> is shown in a seated position. In the seated position, a bearing surface <NUM> on the forward edge 74a of the platform <NUM> abuts the lip <NUM> of the combustor wall <NUM>. The abutment between the lip <NUM> and the bearing surface <NUM> provides a primary seal across the interface between the platform <NUM> and the combustor wall <NUM> to prevent the escape of combustion gases from the core flow path.

The lip <NUM> defines a first annular slot <NUM> and the forward edges 74a of the platforms <NUM> of the vanes <NUM> collectively define a second annular slot <NUM>. The first and second annular slots <NUM>/<NUM> together define an annular seal slot <NUM>. An annular feather seal <NUM> is entrapped in the annular seal slot <NUM> between the combustor wall <NUM> and the platform <NUM>. As will be described further below, the platform <NUM> can move axially away from the lip <NUM>, thereby opening a gap through which combustion gases can escape. In this regard, the annular feather seal <NUM> serves as a secondary seal across the interface between the platform <NUM> and the combustor wall <NUM> to prevent the escape of combustion gases from the core gaspath.

As shown in <FIG> and represented at <NUM>, the vanes <NUM> have rotational play about their radial axes A1 under aerodynamic loads. For instance, although the vanes <NUM> are generally statically mounted, due to tolerances in manufacturing and assembly, the vanes <NUM> can shift somewhat from their proper design positions in which the bearing surfaces <NUM> are seated against the lip <NUM>. Flow of combustion gases from the combustor <NUM> impinges against the second side 80d of the airfoil section <NUM>, particularly toward the trailing end 80b, thereby generating a rotational force about the axis A1. In combination with the play in the position of the vanes <NUM>, the rotational forces can cause the vanes <NUM> to rotate from their design positions. The rotation tends to shift one side of the vanes <NUM> in an axially aft direction, unseating the bearing surface <NUM> from the lip <NUM>. In the unseated position there is thus a divergent gap <NUM> across the interface between the platform <NUM> and the combustor wall <NUM> through which combustion gases can escape. That is, one corner of the platform <NUM> at the forward edge 74a remains in contact with, or at least in close proximity to, the lip <NUM>, while the opposite corner of the platform <NUM> at the forward edge 74a shifts axially aftwards.

In particular, the divergent gap <NUM> presents an unusual sealing challenge because the vanes <NUM> may dynamically move between the seated and unseated positions during engine operation and the forward edges 74a of the platforms <NUM> and the lip <NUM> are non-parallel when in the unseated position. Thus, seals that cannot accommodate dynamic movement or seals that rely on parallel sides may not provide a desired level of sealing. In this regard, the annular feather seal <NUM> is able to address both the dynamic movement and the non-parallel nature of the divergent gap <NUM>.

For example, the configuration of the feather seal <NUM> and the seal slot <NUM> facilitate dynamic sealing of the divergent gap <NUM>. In one example, the seal slot <NUM> defines a slot radial thickness t1 and the feather seal <NUM> defines a seal radial thickness t2, where t2 is less than t1. The seal slot <NUM> also defines a slot axial width w1 and the feather seal <NUM> defines a seal axial width w2, where w2 is less than w1. That is, the feather seal <NUM> is smaller in cross-section than the seal slot <NUM>. This permits the feather seal <NUM> to shift dynamically within the seal slot <NUM> to accommodate shifts in the position of the vanes <NUM>. Additionally, the seal axial width w2 is larger (i.e., wider) than the divergent gap <NUM>, to maintain sealing when the vane <NUM> is in the unseated state. For instance, the play in the vanes <NUM> may be determined or estimated during engine design to determine or estimate the maximum size of the divergent gap <NUM>. The seal axial width w2 is then selected to be larger than the maximum size in order to ensure sealing entirely along the divergent gap <NUM>.

The first and second annular slots <NUM>/<NUM> that define the annular seal slot <NUM> are also configured to bias the feather seal <NUM> to a sealed position. For example, the first slot <NUM> is defined by radially inner and outer walls 82a/82b of the combustor wall <NUM>, and the second slot <NUM> is defined by radially inner and outer walls 84a/84b of the platform <NUM>. The outer wall 82b and the outer wall 84b abut (at bearing surface <NUM> and lip <NUM>). The inner wall 82a and the inner wall 84a are axially spaced apart such that there is a gap <NUM> there between. There is a high pressure region "P" (<FIG>) radially inwards of the combustor wall <NUM> and the platform <NUM>. The pressure in the high pressure region is greater than the pressure in the core gaspath. The high pressure communicates through the gap <NUM> into the seal slot <NUM> to bias the feather seal <NUM> radially outwards, against the outer walls 82b/84b, thereby maintaining the feather seal <NUM> in a sealed position.

To further permit communication of the high pressure in addition to the gap <NUM>, and also reduce weight, at least one of the inner wall 82a or the inner wall 84a is scalloped. <FIG> shows an example of the inner wall 82a, although it is to be understood that the example is also applicable to the inner wall 84a. As shown, the inner wall 82a includes tabs <NUM> and axial slots <NUM> that are circumferentially between the tabs <NUM> to provide a scalloped configuration. The axial slots <NUM> provide additional area for communication of the high pressure into the seal slot <NUM> to bias the feather seal <NUM> radially outwards.

The feather seal <NUM> is also configured to dynamically adapt in diametric size to maintain sealing. <FIG> shows an isolated axial view of the feather seal <NUM>. The feather seal is a split ring. The split ring has first and second ends 88a/88b that are overlapping. The ends 88a/88b are thus free to move relative to one another. The split ring can thereby readily expand and contract in diameter. For instance, the split ring may thermally expand and contract with thermal transients in the engine <NUM>. If constrained, as an endless ring, such expansion or contraction may cause a seal to unseat from its sealing position. However, because the ends 88a/88b can move, the feather seal <NUM> can readily expand and contract and thereby maintain a sealing position against the radially outer walls 82b/84b in the seal slot <NUM> under various thermal conditions.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a combustor (<NUM>) disposed about an engine central axis (A) and including a combustor wall (<NUM>, <NUM>) and a combustion chamber (<NUM>), the combustor wall (<NUM>, <NUM>) having a lip (<NUM>) at an exit region (<NUM>) of the combustion chamber (<NUM>);
a circumferential row of vanes (<NUM>) adj acent the exit region (<NUM>), each said vane (<NUM>) including a platform (<NUM>, <NUM>) and an airfoil section (<NUM>) extending from the platform (<NUM>, <NUM>), the platform (<NUM>, <NUM>) defining forward and trailing edges (74a, 74b) and first and second circumferential side edges (74c, 74d) joining the forward and trailing edges (74a, 74b), the forward edge (74a) being adj acent the lip (<NUM>) of the combustor wall (<NUM>, <NUM>), wherein the lip (<NUM>) defines a first annular slot (<NUM>) and the forward edges (74a) collectively define a second annular slot (<NUM>), the first and second annular slots (<NUM>, <NUM>) together defining an annular seal slot (<NUM>) and the first annular slot (<NUM>) is defined by radially inner and outer walls (82a, 82b) of the combustor wall (<NUM>, <NUM>) and the second annular slot (<NUM>) is defined by radially inner and outer walls (84a, 84b) of the platform (<NUM>, <NUM>),
characterised in that:
an annular feather seal (<NUM>) is entrapped in the annular seal slot (<NUM>) between the combustor wall (<NUM>, <NUM>) and the platform (<NUM>, <NUM>);
the annular feather seal (<NUM>) is a split ring, the split ring having overlapping ends (88a, 88b); and
at least one of the radially inner wall (82a) of the first annular slot (<NUM>) or the radially inner wall (84a) of the second annular slot (<NUM>) is scalloped.