METHODS AND APPARATUS FOR ASSEMBLING TURBINE ENGINES

Abstract

A method for assembling a gas turbine engine includes providing a turbine nozzle including an inner band, an outer band, and at least one vane extending between the inner and outer bands, coupling the turbine nozzle within the gas turbine engine, such that the turbine nozzle is downstream from a combustor, and coupling the turbine nozzle to the combustor using at least a first annular interface seal, wherein the first annular interface seal substantially seals a gap defined between the turbine nozzle and the combustor, and wherein the first annular interface seal includes at least a first ply and a second ply that is coupled to the first ply.

Description

BACKGROUND OF THE INVENTION

This invention relates generally to turbine engines and more particularly, to methods and apparatus for assembling gas turbine engines.

Known gas turbine engines include combustors which ignite fuel-air mixtures which are then channeled through a turbine nozzle assembly towards a turbine. At least some known turbine nozzle assemblies include a plurality of arcuate nozzle segments arranged circumferentially. At least some known turbine nozzles include a plurality of circumferentially-spaced hollow airfoil vanes coupled by integrally-formed inner and outer band platforms. More specifically, the inner band forms a portion of the radially inner flowpath boundary and the outer band forms a portion of the radially outer flowpath boundary.

Within known engine assemblies, an interface defined between the turbine nozzle and an aft end of the combustor is known as a fish-mouth seal. More specifically, within such engine assemblies, leading edges of the turbine nozzle outer and inner band platforms are generally aligned upstream with respect to a leading edge of each airfoil vane extending therebetween. Accordingly, in such engine assemblies, when hot combustion gases discharged from the combustor approach the nozzle vane leading edge, a pressure or bow wave reflects from the vane leading edge stagnation and propagates a distance upstream from the nozzle assembly, causing circumferential pressure variations across the band leading edges and a non-uniform gas pressure distribution. The pressure variations may cause localized nozzle oxidation and/or localized high temperature gas injection, each of which may decrease engine efficiency. Moreover, such pressure variations may also cause the vane leading edge to operate at an increased temperature in comparison to the remainder of the vane.

BRIEF SUMMARY OF THE INVENTION

In one aspect, a method for assembling a gas turbine engine is provided. The method includes providing a turbine nozzle including an inner band, an outer band, and at least one vane extending between the inner and outer bands, coupling the turbine nozzle within the gas turbine engine, such that the turbine nozzle is downstream from a combustor, and coupling the turbine nozzle to the combustor using at least a first annular interface seal, wherein the first annular interface seal substantially seals a gap defined between the turbine nozzle and the combustor, and wherein the first annular interface seal includes at least a first ply and a second ply that is coupled to the first ply.

In another aspect, a turbine engine nozzle assembly for use in a turbine engine including a combustor is provided. The engine nozzle assembly includes an outer band comprising a leading edge, a trailing edge, and a body extending therebetween, an inner band comprising a leading edge, a trailing edge, and a body extending therebetween, at least one vane extending between said outer and inner bands, said at least one vane comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge, said at least one vane leading edge positioned downstream from said inner and outer band leading edges, and a system configured to couple said nozzle to the combustor, said system comprising a first annular interface seal coupled between said outer band and the combustor and a second annular interface seal coupled between said inner band and the combustor, said first and second annular interface seals each comprising a first ply and a second ply that is coupled to said first ply.

In a further aspect, a gas turbine engine assembly is provided. The gas turbine engine assembly includes a combustor, a turbine nozzle assembly downstream from and in flow communication with said combustor, said nozzle assembly comprising an outer band, an inner band, at least one vane extending between said outer and inner bands, and a leading edge fillet, said at least one vane comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge, and a system configured to couple said nozzle to the combustor, said system comprising a first annular interface seal coupled between said outer band and the combustor and a second annular interface seal coupled between said inner band and the combustor, said first and second annular interface seals each comprising a first ply and a second ply that is coupled to said first ply.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of an exemplary gas turbine engine;

FIG. 2 is a side view of an exemplary turbine nozzle that may be used with the gas turbine engine shown in FIG. 1;

FIG. 3 is a perspective view of the turbine nozzle shown in FIG. 2;

FIG. 4 is an enlarged side view of an exemplary interface seal that may be used with the turbine nozzle shown in FIGS. 2 and 3;

FIG. 5 is a side view of the turbine nozzle shown in FIGS. 2 and 3 coupled to a combustor that may be used with the engine shown in FIG. 1 with the interface seal shown in FIG. 4;

FIG. 6 illustrates the turbine nozzle and interface seal shown in FIG. 5 and a first wear coating;

FIG. 7 illustrates the turbine nozzle and interface seal shown in FIG. 5 and a second wear coating;

FIG. 8 illustrates the turbine nozzle and interface seal shown in FIG. 5 and a third wear coating;

FIG. 9 illustrates the turbine nozzle and interface seal shown in FIG. 5 and a fourth wear coating.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a schematic illustration of an exemplary gas turbine engine 10 including a fan assembly 12, a high pressure compressor 14, and a combustor 16. Engine 10 also includes a high pressure turbine 18 and a low pressure turbine 20. Fan assembly 12 and turbine 20 are coupled by a first shaft 21, and compressor 14 and turbine 18 are coupled by a second shaft 22. In one embodiment, gas turbine engine 10 is an LM2500 engine commercially available from General Electric Aircraft Engines, Cincinnati, Ohio. In another embodiment, gas turbine engine 10 is a CFM engine commercially available from General Electric Aircraft Engines, Cincinnati, Ohio.

In operation, air flows through fan assembly 12 supplying compressed air to high pressure compressor 14. The highly compressed air is delivered to combustor 16. Airflow from combustor 16 is channeled through a turbine nozzle (not shown in FIG. 1) to drive turbines 18 and 20, prior to exiting gas turbine engine 10 through an exhaust nozzle 24.

FIG. 2 is a side view of an exemplary turbine nozzle 50 that may be used with a gas turbine engine, such as turbine engine 10 (shown in FIG. 1). FIG. 3 is a perspective view of turbine nozzle 50. In the exemplary embodiment, nozzle 50 is one segment of a plurality of segments that are positioned circumferentially to form a nozzle assembly (not shown) within the gas turbine engine. Nozzle 50 includes at least one airfoil vane 52 that extends between an arcuate radially outer band or platform 54, and an arcuate radially inner band or platform 56. More specifically, in the exemplary embodiment, outer band 54 and the inner band 56 are each integrally-formed with airfoil vane 52.

Vane 52 includes a pressure-side sidewall 60 and a suction-side sidewall 62 that are connected at a leading edge 64 and at an chordwise-spaced trailing edge 66 such that a cooling cavity 68 is defined between sidewalls 60 and 62. Vane sidewalls 60 and 62 each extend radially between bands 54 and 56 and in the exemplary embodiment, sidewall 60 is generally concave, and sidewall 62 is generally convex.

Outer and inner bands 54 and 56 each include a leading edge 70 and 72, respectively, a trailing edge 74 and 76, respectively, and a platform body 78 and 80, respectively, extending therebetween. Airfoil vane(s) 52 are oriented such that outer and inner band leading edges 70 and 72, respectively, are each a distance d upstream from airfoil vane leading edge 64. Distance d is variably selected to ensure that leading edges 70 and 72 are upstream from vane leading edge 64, and to facilitate bands 54 and 56 preventing hot gas injections along vane leading edge 64, as described in more detail below.

In the exemplary embodiment, inner band 56 includes an aft flange 90 that extends radially inwardly therefrom. More specifically, flange 90 extends radially inwardly from band 56 with respect to a radially inner surface 92 of band 56. Inner band 56 also includes a forward flange 94 that extends radially inward therefrom. Forward flange 94 is positioned between inner band leading edge 72 and aft flange 90, and extends radially inwardly from band 56. In the exemplary embodiment, an upstream side 100 of forward flange 94 is substantially planar between a radially outermost surface 102 of flange 94 and radially inner surface 92.

Inner band 56 also includes a plurality of circumferentially-spaced radial tabs 110 that extend radially inwardly therefrom. More specifically, in the exemplary embodiment, the number of radial tabs 110 is the same as the number of vanes 52. In the exemplary embodiment, each tab 110 includes a substantially parallel upstream and downstream surfaces 120 and 122, respectively. Radial tabs 110 are spaced a distance d₂ downstream from forward flange 94 such that a retention channel 130 is defined between each radial tab 110 and forward flange 94.

In the exemplary embodiment, outer band 54 includes an aft flange 140 that extends generally radially outwardly therefrom. More specifically, flange 140 extends radially outwardly from band 54 with respect to a radially outer surface 142 of band 54. Outer band 54 also includes a forward flange 144 that extends radially outward therefrom. Forward flange 144 is positioned between outer band leading edge 70 and aft flange 140, and extends radially outwardly from band 54. In the exemplary embodiment, an upstream side 146 of forward flange 144 is substantially planar between a radially outermost surface 147 of flange 144 and radially outer surface 142.

Outer band 54 also includes a plurality of circumferentially-spaced radial tabs 160 that extend radially outwardly therefrom. More specifically, in the exemplary embodiment, the number of radial tabs 160 is the same as the number of vanes 52. In the exemplary embodiment, each tab 160 includes substantially parallel upstream and downstream surfaces 162 and 164, respectively. Radial tabs 160 are spaced a distance d₃ downstream from forward flange 144 such that a retention channel 166 is defined between each radial tab 160 and forward flange 144. In the exemplary embodiment, channels 166 are approximately the same size as channels 130.

Turbine nozzle 50 also includes a plurality of leading edge fillets 170. Fillets 170 are generally larger than fillets used with known turbine nozzles and extend between outer platform 54 and vane 52 in a tip area 180 of each vane leading edge 64, and between inner platform 56 and vane 52 in a hub area 182 of each vane leading edge 64. Specifically, within tip area 180, fillets 170 are blended from vane leading edge 64 across a radially inner surface 184 of outer platform 54 and towards outer band leading edge 70. Moreover, within hub area 182, fillets 170 are blended from vane leading edge 64 across a radially outer surface 186 of inner platform 56 and towards inner band leading edge 72. Accordingly, nozzle vane leading edge 64 is enlarged within both hub area 182 and tip area 180 such that fillets 170 facilitate accelerating the flow passing thereby.

In the exemplary embodiment, fillets 170 are formed with a plurality of cooling openings 190 that extend through fillets 170 and are configured to discharge cooling air inwardly into the boundary flow flowing over vane 52. Specifically, each cooling opening 190 is oriented towards a pitch-line of vane 52 and such that openings 190 facilitate energizing the flow momentum in the boundary layer, such that the formation of horseshoe vortices upstream from leading edge 64 is facilitated to be reduced. The reduction in the formation of the horseshoe vortices facilities improving aerodynamic efficiency. Moreover, the plurality of cooling openings 190 also facilitate reducing surface heating and an operating temperature of vane 52.

During operation, the location of inner and outer bands 56 and 54, respectively, with respect to vane leading edge 64 facilitates reducing hot gas injections along vane leading edge 64. Rather, the combination of enlarged fillets 170 and cooling holes 190 facilitates accelerating the flow and energizing the flow momentum in the boundary layer, such that the formation of horseshoe vortices are facilitated to be reduced. As a result, aerodynamic efficiency is facilitated to be improved and the operating temperature of nozzle airfoil vane 52 is facilitated to be reduced. As such, a useful life of turbine nozzle 50 is facilitated to be extended.

FIG. 4 is an enlarged side view of an exemplary interface seal 200 that may be used with turbine nozzle 50 (shown in FIGS. 2 and 3). In the exemplary embodiment, interface seal 200 is a substantially L-shaped seal having a length 204 and a width 208 that is less than length 204. Interface seal 200 is fabricated to include a plurality of plies of material. In the exemplary embodiment, interface seal 200 includes at least a first ply 220 and a second ply 222 that are coupled together to form a two-ply interface seal 200. Interface seal 200 also has an upstream end 224 and a downstream end 226. In the exemplary embodiment, the first and second plies 220 and 222 are each fabricated from a metallic material. For example, the first and second plies 220 and 222 may be fabricated utilizing a nickel based alloy such as HS 188.

During assembly, first and second plies 220 and 222 are coupled together at at least one discrete location to facilitate reducing or damping vibrations transmitted therethrough. For example, first and second plies 220 and 222 may be spot welded at a point 230 that is located approximately at a midpoint along first portion 202. Optionally, first and second plies 220 and 222 may be coupled together at a plurality of discrete locations using various techniques to improve damping capability. In the exemplary embodiment, interface seal 200 has a width or thickness 232 that is sized to downstream end 226 to be at least partially inserted into either channel 166 or channel 130 (shown in FIGS. 2 and 3). Accordingly, downstream end 226 is sized for insertion within retention channels 166 and 130.

As discussed above, interface seal 200 is fabricated from a resilient metallic material that resists deformation. For example, interface seal 200 may be fabricated utilizing a shape memory material. In a further alternative embodiment, interface seal 200 is fabricated from any material that enables interface seal 200 to function as described herein.

FIG. 5 is a side view of turbine nozzle 50 coupled to combustor 16 using two interface seals 200. Combustor 16 includes a combustion zone 240 that is formed by annular, radially inner and radially outer supporting members 244 and 246, respectively, and combustor liners 250. Combustor liners 250 shield the outer and inner supporting members from heat generated within combustion zone 240. More specifically, combustor 16 includes an annular inner liner 256 and an annular outer liner 258. Liners 256 and 258 define combustion zone 240 such that combustion zone 240 extends from a dome assembly (not shown) downstream to turbine nozzle 50. Outer and inner liners 258 and 256, respectively each include a plurality of separate panels 260 which include a series of steps 262, each of which form a distinct portion of combustor liners 250.

Each liner 256 and 258 also includes an annular support flange, or aft flange, 270 and 272, respectively. Specifically, each support flange 270 and 272 couples an aft end 274 and 276 of each respective liner 256 and 258 to supporting members 244 and 246. More specifically, the coupling of each support flange 270 and 272 to each supporting member 244 and 246 forms an annular gap or fishmouth opening 278.

Each support flange 270 and 272 includes a radial portion 280 and a conical datum area 282. Each radial portion 280 is formed with a plurality of preferential cooling openings or jets 284 that extend therethrough to facilitate discharging cooling air towards nozzle 50. Air discharged from jets 284 facilitates reducing the formation of horseshoe vortices upstream from vane leading edge 64 and thus facilitates improving aerodynamic efficiency of nozzle 50. Each conical datum area 282 extends integrally outward and upstream from each radial portion 280 such that conical datum area 282 defines a radially inner portion 286 of each fishmouth opening 278. A radial outer portion 288 of each fishmouth opening 278 is defined by each supporting member 244 or 246. Fishmouth opening 278 is used to couple each interface seal 200 between combustor 16 and nozzle 50.

During assembly, combustor 16 is coupled to nozzle 50 utilizing a system that includes a plurality of interface seals 200. More specifically, the system includes at least a first annular interface seal 201, and a second annular interface seal 203. First and second annular interface seals 201 and 203 are each implemented utilizing interface seal 200 shown in FIG. 4. In the exemplary embodiment, the upstream end 224 of first and second interface seals 201 and 203 are each securely coupled within a respective fishmouth opening 278, and the downstream ends 226 are each inserted within a respective retention channel 130 and 166.

When the engine is fully assembled, interface seals 201 and 203 couple combustor 16 to nozzle 50 and also facilitate sealing between combustor 16 and nozzles 50. As such, a mechanically flexible seal arrangement is provided which provides structural stability and support to the aft end of combustor 16. Moreover, the assembly of interface seals 201 ad 203 between combustor 16 and nozzle 50 is generally less labor intensive and less time-consuming than the assembly of known seal interfaces used with other gas turbine engines.

In the exemplary embodiment, at least one of nozzle 50, interface seal 201, and/or interface seal 203 may be at least partially coated with a wear coating material to reduce wear between nozzle 50 and the respective interface seals, and thus increase the useful life of both the turbine nozzle and interface seals. Moreover, although the wear coatings described herein are illustrated with respect to interface seal 201 and retention channel 166 defined between forward flange 144 and tabs 160, it should be realized that the wear coatings may also be applied to interface seal 201 and retention channel 130 defined between forward flange 94 and tabs 110.

For example, FIG. 6 illustrates turbine nozzle 50 and interface seal 201 including a wear coating 300 that is formed on downstream side 148 of flange 144. FIG. 7 illustrates turbine nozzle 50 and interface seal 201 including wear coating 300 that is applied to downstream side 148 of flange 144 and a wear coating 302 that is applied upstream side 162 of tabs 160. FIG. 8 illustrates turbine nozzle 50 and interface seal 201 including wear coatings 300, 302, and a wear coating 304 that is applied to an upstream side 400 of interface seal 201 proximate to downstream end 206. FIG. 9 illustrates turbine nozzle 50 and interface seal 201 including wear coatings 300, 302, 304, and a wear coating 306 that is applied to a downstream side 402 of interface seal 201. In the exemplary embodiment, wear coatings 300, 302, 304, and 306 are each a brazed T800 wear coating, or optionally, a braze tape that is applied to the surfaces described herein.

In each embodiment, the above-described turbine nozzles include an inner band and an outer band that each extend upstream a distance from the vane leading edge to facilitate reducing hot gas injection along the vane leading edge. Moreover, because each inner and outer band extends upstream from the vane leading edge, each band accommodates enlarged fillets in comparison to known turbine nozzles. The combination of the inner and outer bands, the impingement jets extending through the combustor support flanges, and the cooling openings extending through the fillets facilitates reducing an operating temperature of the nozzle vanes, reducing the formation of horseshoe vortices upstream from each vane leading edge, and improving the aerodynamic efficiency of the nozzle. Moreover, the interface seals are fabricated utilizing a multiply metallic material and extend between the combustor and the turbine nozzle provide increased structural support to the combustor while being biased in a sealing arrangement with the turbine nozzle. As a result, a useful life of the turbine nozzle is facilitated to be extended in a reliable and cost effective manner.

Exemplary embodiments of turbine nozzles are described above in detail. The interface seals, fillets, and cooling openings and jets are not limited to use with the specific nozzle embodiments described herein, but rather, the such components can be utilized independently and separately from other turbine nozzle components described herein. Moreover, the invention is not limited to the embodiments of the nozzle assemblies described above in detail. Rather, other variations of nozzles assembly embodiments may be utilized within the spirit and scope of the claims.

While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.

Claims

1. A method for assembling a gas turbine engine, said method comprising: providing a turbine nozzle including an inner band, an outer band, and at least one vane extending between the inner and outer bands; coupling the turbine nozzle within the gas turbine engine, such that the turbine nozzle is downstream from a combustor; and coupling the turbine nozzle to the combustor using at least a first annular interface seal, wherein the first annular interface seal substantially seals a gap defined between the turbine nozzle and the combustor, and wherein the first annular interface seal includes at least a first ply and a second ply that is coupled to the first ply.

2. A method in accordance with claim 1 further comprising coupling the turbine nozzle to the combustor using at least a first annular interface seal that includes at least a first metallic ply and a second metallic ply that is coupled to the first metallic ply.

3. A method in accordance with claim 1 further comprising coupling the turbine nozzle to the combustor using at least a first annular interface seal that includes at least a first nickel based alloy ply and a second nickel based alloy ply that is coupled to the first nickel based alloy ply.

4. A method in accordance with claim 1 further comprising applying a wear coating to at least a portion of the first annular interface seal.

5. A method in accordance with claim 1 further comprising applying a wear coating to at least a portion of the turbine nozzle.

6. A method in accordance with claim 1 further comprising applying a wear coating to at least a portion of the first annular interface seal and the turbine nozzle.

7. A turbine engine nozzle assembly for use in a turbine engine including a combustor, said engine nozzle assembly comprising: an outer band comprising a leading edge, a trailing edge, and a body extending therebetween; an inner band comprising a leading edge, a trailing edge, and a body extending therebetween; at least one vane extending between said outer and inner bands, said at least one vane comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge, said at least one vane leading edge positioned downstream from said inner and outer band leading edges; and a system configured to couple said nozzle to the combustor, said system comprising a first annular interface seal coupled between said outer band and the combustor and a second annular interface seal coupled between said inner band and the combustor, said first and second annular interface seals each comprising a first ply and a second ply that is coupled to said first ply.

8. A turbine engine nozzle assembly in accordance with claim 7 wherein at least one of said first annular interface seal and said second annular interface seal comprises an L-shaped annular body.

9. A turbine engine nozzle assembly in accordance with claim 7 wherein said first and second annular interface seals each comprise a first metallic ply and a second metallic ply that is coupled to said first metallic ply.

10. A turbine engine nozzle assembly in accordance with claim 7 wherein said first and second annular interface seals each comprise a first nickel based alloy ply and a second nickel based alloy ply that is coupled to said first nickel based alloy ply.

11. A turbine engine nozzle assembly in accordance with claim 7 further comprising a wear coating applied to at least one of said first and second annular interface seals.

12. A turbine engine nozzle assembly in accordance with claim 7 further comprising a wear coating applied to at least a portion of said turbine nozzle.

13. A turbine engine nozzle assembly in accordance with claim 7 further comprising a wear coating applied to at least one of said first and second annular interface seals and to at least a portion of said turbine nozzle.

14. A gas turbine engine comprising: a combustor; a turbine nozzle assembly downstream from and in flow communication with said combustor, said nozzle assembly comprising an outer band, an inner band, at least one vane extending between said outer and inner bands, and a leading edge fillet, said at least one vane comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge; and a system configured to couple said nozzle to the combustor, said system comprising a first annular interface seal coupled between said outer band and the combustor and a second annular interface seal coupled between said inner band and the combustor, said first and second annular interface seals each comprising a first ply and a second ply that is coupled to said first ply.

15. A gas turbine engine in accordance with claim 14 wherein at least one of said first annular interface seal and said second annular interface seal comprises an L-shaped annular body.

16. A gas turbine engine in accordance with claim 14 wherein said first and second annular interface seals each comprise a first metallic ply and a second metallic ply that is coupled to said first metallic ply.

17. A gas turbine engine in accordance with claim 14 wherein said first and second annular interface seals each comprise a first nickel based alloy ply and a second nickel based alloy ply that is coupled to said first nickel based alloy ply.

18. A gas turbine engine in accordance with claim 14 further comprising a wear coating applied to at least one of said first and second annular interface seals.

19. A gas turbine engine in accordance with claim 14 further comprising a wear coating applied to at least a portion of said turbine nozzle.

20. A gas turbine engine in accordance with claim 14 further comprising a wear coating applied to at least one of said first and second annular interface seals and to at least a portion of said turbine nozzle.