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.

Gas turbine stator vane assemblies typically include a plurality of vane segments which collectively form the annular vane assembly. Each vane segment includes one or more airfoils extending between an outer platform and an inner platform. The inner and outer platforms collectively provide radial boundaries to guide core gas flow past the airfoils. Core gas flow may be defined as gas exiting the compressor passing directly through the combustor and entering the turbine.

Vane support rings support and position each vane segment radially inside of the engine diffuser case. In most instances, cooling air bled off of the fan is directed into an annular region between the diffuser case and an outer case, and a percentage of compressor air is directed in the annular region between the outer platforms and the diffuser case, and the annular region radially inside of the inner platforms.

The fan air is at a lower temperature than the compressor air, and consequently cools the diffuser case and the compressor air enclosed therein. The compressor air is at a higher pressure and lower temperature than the core gas flow which passes on to the turbine. The higher pressure compressor air prevents the hot core gas flow from escaping the core gas flow path between the platforms. The lower temperature of the compressor flow keeps the annular regions radially inside and outside of the vane segments cool relative to the core gas flow.

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

<CIT> discloses a vane for a gas turbine engine comprising a first airfoil, a first chordal seal located adjacent a first end of the first airfoil, wherein the first chordal seal is located on a rail located on an opposite side of a first platform from the first airfoil and a first pair of transition regions which extend along a pair of edges of the first chordal seal, a second chordal seal located adjacent a second end of the first airfoil, wherein the first chordal seal includes a first edge parallel to a first edge on the second chordal seal and a cusp of material spaced radially inward from the first chordal seal.

<CIT> discloses a prior art nozzle guide vane assembly.

According to the invention, there is provided a vane for a gas turbine engine as set forth in claim <NUM>.

In an embodiment of the above, the first chordal seal includes a second edge parallel to a second edge on the second chordal seal.

In a further embodiment of any of the above, a second pair of transition regions extends along a pair of edges of the second chordal seal.

In a further embodiment of any of the above, there is a second airfoil. The first airfoil and the second airfoil extend between the first platform located at a first end of the first and second airfoils. A second platform is located at a second end of the first and second airfoils.

According to the invention, there is further provided a method of forming a vane for a gas turbine engine as set forth in claim <NUM>.

In an embodiment of the above, a second edge of the first chordal seal adjacent the first end of the airfoil is machined while the component is attached to the fixture. A second edge of the second chordal seal adjacent the second end of the airfoil is machined while the component is attached to the fixture.

In a further example, the engine <NUM> bypass ratio is greater than about six (<NUM>:<NUM>), with an example embodiment being greater than about ten (<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 five. The geared architecture <NUM> may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>.

The fan section <NUM> of the engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach and about <NUM>,<NUM> (<NUM>,<NUM> feet). The flight condition of <NUM> Mach and <NUM>,<NUM> (<NUM>,<NUM> ft), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about <NUM>/s (<NUM> ft / second).

The example gas turbine engine includes fan <NUM> that comprises in one non-limiting embodiment less than about twenty-six fan blades. In another non-limiting embodiment, fan section <NUM> includes less than about twenty fan blades. Moreover, in one disclosed embodiment low pressure turbine <NUM> includes no more than about six turbine rotors. In another non-limiting example embodiment low pressure turbine <NUM> includes about three turbine rotors. A ratio between number of fan blades <NUM> and the number of low pressure turbine rotors is between about <NUM> and about <NUM>. The example low pressure turbine <NUM> provides the driving power to rotate fan section <NUM> and therefore the relationship between the number of turbine rotors in low pressure turbine <NUM> and number of blades <NUM> in fan section <NUM> disclose an example gas turbine engine <NUM> with increased power transfer efficiency.

<FIG> illustrates an enlarged schematic view of the high pressure turbine <NUM>, however, other sections of the gas turbine engine <NUM> could benefit from this disclosure. In the illustrated example, the high pressure turbine <NUM> includes a one-stage turbine section with a first rotor assembly <NUM>. In another example, the high pressure turbine <NUM> could include a two-stage high pressure turbine section.

The first rotor assembly <NUM> includes a first array of rotor blades <NUM> circumferentially spaced around a first disk <NUM>. Each of the first array of rotor blades <NUM> includes a first root portion <NUM>, a first platform <NUM>, and a first airfoil <NUM>. Each of the first root portions <NUM> is received within a respective first rim <NUM> of the first disk <NUM>. The first airfoil <NUM> extends radially outward toward a first blade outer air seal (BOAS) assembly <NUM>.

The first array of rotor blades <NUM> are disposed in the core flow path that is pressurized in the compressor section <NUM> then heated to a working temperature in the combustor section <NUM>. The first platform <NUM> separates a gas path side inclusive of the first airfoils <NUM> and a non-gas path side inclusive of the first root portion <NUM>.

An array of vanes <NUM> are located axially upstream of the first array of rotor blades <NUM>. Each of the array of vanes <NUM> include at least one airfoil <NUM> that extend between a respective vane inner platform <NUM> and an vane outer platform <NUM>. In another example, each of the array of vanes <NUM> include at least two airfoils <NUM> forming a vane double. The vane outer platform <NUM> of the vane <NUM> may at least partially engage the BOAS <NUM>.

As shown in <FIG>, the vane <NUM> includes a first chordal seal <NUM> and a second chordal seal <NUM> on an axially downstream end of the vane <NUM>. The first and second chordal seals <NUM>, <NUM> are inner and outer chordal seals respectively. In this disclosure, axial or axially extending is in relation to the axis A of the gas turbine engine <NUM>. The outer chordal seal <NUM> creates a seal between the vane <NUM> and the BOAS <NUM>. The outer chordal seal <NUM> extends in a chordal direction along an axially facing surface <NUM> of an outer rail <NUM>. The outer rail <NUM> extends radially outward from the vane outer platform <NUM>. By having the outer chordal seal <NUM> extend in the chordal direction, the outer chordal seal <NUM> will be straight and extend between opposing circumferential ends of the outer rail <NUM>.

The outer chordal seal <NUM> includes an axially facing surface <NUM> that faces axially downstream relative to the axis A of the gas turbine engine <NUM>. The axially facing surface <NUM> is axially spaced from the axially facing surface <NUM> by a pair of transition regions <NUM>. The pair of transition regions <NUM> includes a pair of fillets having a radius of curvature.

The inner chordal seal <NUM> creates a seal between the vane <NUM> and a portion of the static structure <NUM>. The inner chordal seal <NUM> extends in a chordal direction along an axially facing surface <NUM> of an inner rail <NUM> extending radially inward from the vane inner platform <NUM>. By having the inner chordal seal <NUM> extend in the chordal direction, the inner chordal seal <NUM> will be straight and extend between opposing circumferential ends of the vane inner platform <NUM>.

In the illustrated example, the portion of the static structure <NUM> creating the seal with the inner chordal seal <NUM> is a flange <NUM> on a tangent on board injector (TOBI). However, another portion of the static structure <NUM> could be used to engage the inner chordal seal <NUM>.

The inner chordal seal <NUM> includes an axially facing surface <NUM> that faces axially downstream relative to the axis A of the gas turbine engine <NUM>. The axially facing surface <NUM> is spaced from the axially facing surface <NUM> by a pair of transition regions <NUM>. The pair of transition regions <NUM> includes a pair of fillets having a radius of curvature.

As shown in <FIG>, a cusp <NUM> is located on a radially inner portion of the inner rail <NUM>. The cusp <NUM> is at least partially defined by one of the transition regions <NUM> along an axially downstream edge and by a recess <NUM> along an axially forward edge. In the illustrated example, the recess <NUM> includes a pair of angled surfaces. In another example, the recess <NUM> could include a fillet having a radius of curvature.

Axial positions of the outer chordal seal <NUM> and the inner chordal seal <NUM> may vary slightly from one another due to manufacturing tolerances and nominal dimensions of the vane <NUM> in a cold state. Because of the variations in the vane <NUM>, corresponding pairs of edges on the outer chordal seal <NUM> and inner chordal seal <NUM> would engage the BOAS <NUM> and the flange <NUM>, respectively, and form the seal.

In one example, when the vane outer platform <NUM> is shifted axially rearward of the vane inner platform <NUM>, a first edge 100a of the outer chordal seal <NUM> engages the BOAS <NUM> and a first edge 102a of the inner chordal seal <NUM> engages the flange <NUM>. In another example, when the vane outer platform <NUM> is shifted axially forward of the vane inner platform <NUM>, a second edge 100b of the outer chordal seal <NUM> engages the BOAS <NUM> and a second edge 102b of the inner chordal seal <NUM> engages the flange <NUM>. The first edges 100a, 102a are located on a radially outer side of the outer chordal seal <NUM> and the inner chordal seal <NUM>, respectively, and the second edges 100b, 102b are located on a radially inner side of the outer chordal seal <NUM> and the inner chordal seal <NUM>, respectively.

In order to improve the effectiveness of the outer and inner chordal seals <NUM> and <NUM>, the first edge 100a must be parallel to the first edge 102a and the second edge 100b must be parallel to the second edge 102b. By improving the parallelism between the corresponding edges on the outer and inner chordal seals <NUM>, <NUM>, the corresponding edges are able to maintain a line of contact with the BOAS <NUM> and static structure <NUM>, respectively, when the deflection between the static structure <NUM> attached to the vane outer platform <NUM> and the static structure <NUM> attached to inner platform <NUM> varies.

In order to improve the parallelism and simplify the manufacturing process of the vane <NUM>, the first edges 100a, 102a and the second edges 100b, 102b are formed during the same machining process. By forming the first edges 100a, 102a and the second edges 100b, 102b in the same jig during machining, variations in parallelism between the first edges 100a, 102a and the second edges 100b, 102b are reduced. The variations in parallelism are reduced because the vane <NUM> does not need to be mounted into a second jig which can reduce parallelism if the vane <NUM> is not aligned perfectly in the second jig.

Claim 1:
A vane (<NUM>) for a gas turbine engine (<NUM>) comprising:
a first airfoil (<NUM>);
a first chordal seal (<NUM>) located adjacent a first end of the first airfoil (<NUM>), wherein the first chordal seal (<NUM>) is located on a rail (<NUM>) located on an opposite side of a first platform (<NUM>) from the first airfoil (<NUM>) and a first pair of transition regions (<NUM>) extend along a pair of edges of the first chordal seal (<NUM>);
a second chordal seal (<NUM>) located adjacent a second end of the first airfoil (<NUM>), wherein the first chordal seal (<NUM>) includes a first edge (102a) parallel to a first edge (100a) on the second chordal seal (<NUM>); and
a cusp of material (<NUM>) spaced radially inward from the first chordal seal (<NUM>);
characterised by further comprising:
a recess (<NUM>) in the rail (<NUM>) on an opposite side of the cusp of material (<NUM>) from the first chordal seal (<NUM>), wherein the cusp of material (<NUM>) is defined by one of the first pair of transition regions (<NUM>) along an axially downstream edge of the cusp of material (<NUM>) and by the recess (<NUM>) along an axially forward edge of the cusp of material (<NUM>); and
wherein the second chordal seal (<NUM>) extends in a chordal direction along an axially facing surface (<NUM>) of an outer rail (<NUM>), the second chordal seal (<NUM>) includes an axially facing surface (<NUM>) that is configured to face axially downstream relative to an axis (A) of the gas turbine engine (<NUM>), and the axially facing surface (<NUM>) is axially spaced from the axially facing surface (<NUM>) by a second pair of transition regions (<NUM>), wherein the first and second pair of transition regions (<NUM>, <NUM>) each include a pair of fillets having a radius of curvature.