GAS TURBINE AIRFOIL WITH COOLING ENHANCEMENT

Abstract

A turbine nozzle vane segment includes one or more nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a peripheral edge wall extending between a leading edge and a trailing edge of the vane. In one exemplary embodiment, at least one substantially radially-oriented cooling channel is formed in the peripheral edge wall at the leading edge, with openings at opposite ends of the cooling channel. The location and length of the cooling channels may vary about the peripheral edge wall, and the inner cavity of the vane may be provided with ribs extending along and adjacent the one or more cooling channels to reinforce the wall and to also provide additional cooling surface areas in the inner cavity.

Description

BACKGROUND OF THE INVENTION

This present application relates generally to improving the efficiency and/or operation of turbine engines. More specifically, but not by way of limitation, the present application relates to enhancing the cooling of turbine nozzle vanes or blades.

A gas turbine engine typically includes a compressor, one or more combustors, and at least one turbine section. The compressor and turbine section generally include rows of vanes and buckets that are axially stacked in stages. Each stage may include alternating rows of circumferentially-spaced stator vanes, which are stationary, and circumferentially-spaced buckets, that are mounted on a wheel fixed to the turbine rotor. In operation, the rotor blades in the compressor rotate with the rotor to compress a flow of air supplied to the compressor. Most of the compressed air is mixed with gaseous or liquid fuel in the one or more combustors and ignited to provide a stream of hot gases, which is expanded through the turbine section of the engine, causing rotation of the turbine rotor. Thus, the energy contained in the fuel is converted into the mechanical energy of the rotating rotor which may be used to rotate the rotor blades of the compressor such that the supply of compressed air needed for combustion is produced, as well as the coils of a generator, such that electrical power is produced.

During operation, because of the extreme temperatures in the hot-gas path, the velocity of the working fluid, and the rotational velocity of the engine, the rotating buckets (or airfoils) and the stationary stator vanes become highly stressed due to extreme mechanical and thermal loads.

As one of ordinary skill in the art will appreciate, one strategy for alleviating thermal stresses is through cooling the nozzle vanes and/or buckets such that the temperatures experienced by the vanes and/or buckets are lower than that of the hot-gas path. Effective cooling may, for example, allow these hot gas path components to withstand higher firing temperatures, withstand greater thermo-mechanical stresses at high operating temperatures, and/or extend service life, all of which may allow the turbine engine to be more cost-effective and efficient. One way to cool vanes and buckets during operation is through the use of internal cooling passageways or circuits. Generally, this involves passing relatively cool air, which may be supplied by the compressor, through internal cooling circuits within the vanes or airfoils.

There remains a need, however, to provide more effective and more efficient cooling with respect to, for example, leading edges of the stationary nozzle vanes which are exposed to hot combustion gases, particularly in the first turbine stage, where the highest temperatures and thermal stresses are experienced.

BRIEF DESCRIPTION OF THE INVENTION

In one exemplary but nonlimiting embodiment, there is provided a turbine nozzle vane segment comprising: one or more nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a edge wall including leading and trailing edges; and at least one substantially radially-oriented cooling channel formed in the peripheral edge wall with an inlet at one of the inner and outer side walls.

In another nonlimiting aspect, the invention provides a turbine engine comprising a compressor, at least one combustor and at least one turbine stage including a row of stationary nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a peripheral edge wall including leading and trailing edges; and at least one substantially radially-oriented cooling channel formed in the peripheral edge wall with an inlet provided at one of the inner and outer side walls.

It is also a feature of the invention to provide reinforcing ribs in the inner cavity of the nozzle vane, adjacent and at least partially along the channels in the peripheral edge wall. Accordingly, in still another exemplary aspect, there is provided a turbine engine comprising a compressor, at least one combustor and at least one turbine stage including a row of stationary nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a peripheral edge wall including leading and trailing edges; and a plurality of substantially radially-oriented cooling channels formed about the peripheral edge wall including at the leading edge, the plurality of substantially radially-oriented cooling channels extending at least part way along a radial length between the inner and outer side walls and having inlets in or adjacent one of the inner and outer side walls; and a reinforcing rib extending along an inside surface of the peripheral edge wall adjacent each of the plurality of substantially radially-oriented cooling channels.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic diagram of a conventional gas turbine engine;

FIG. 2 is a partial side elevation of a turbine section of a conventional gas turbine engine;

FIG. 3 is a perspective view of a gas turbine nozzle vane set in accordance with an exemplary but nonlimiting embodiment of the invention;

FIG. 4 is a partially sectioned view illustrating a leading edge cooling channel in accordance with the exemplary but nonlimiting embodiment of the invention;

FIGS. 5 and 6 are enlarged details taken from FIG. 4;

FIG. 7 is a partial perspective view illustrating the exemplary cooling channel in accordance with one embodiment;

FIG. 8 is a partial perspective view illustrating the exemplary cooling channel in accordance with another embodiment;

FIG. 9 is a partial perspective view illustrating the exemplary cooling channel in accordance with still another embodiment;

FIG. 10 is a partial top plan view of a nozzle vane provided with cooling channels in accordance with another exemplary but nonlimiting embodiment;

FIG. 11 is a partial section of a nozzle vane having a cooling channel with different diameters in accordance with another exemplary embodiment; and

FIG. 12 is a partial section of a nozzle vane showing an array of cooling channels adjacent a leading edge of the vane in accordance with another exemplary embodiment.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 illustrates in schematic form a gas turbine system 10 that, as described further herein, includes stationary vanes and rotating blades or buckets that may be provided with internal cooling circuits. In this otherwise conventional arrangement, air supplied via inlet 12 is pressurized in a compressor 14 and mixed with fuel in one or more combustors 16 where it is ignited to thereby generate hot combustion gases. Energy is extracted from the combustion gases in one or more turbine stages 18 disposed downstream of the combustors, to drive a generator 20 producing electric power. The extracted energy may also be used to drive the compressor 14, and note that the turbine rotor 22 may be common to the compressor, turbine stages and generator. The invention described herein, however, is not limited to just the illustrated gas turbine system. Further in that regard, the cooling circuits described herein are fully compatible with various film-cooling configurations utilizing air flowing through the cooling circuit passages or cavities.

With additional reference to FIG. 2, in a typical gas turbine configuration, multiple combustors 16 may be annularly disposed about the turbine rotor, with multiple transition pieces 24 that direct the hot combustion gases from the respective combustors to the gas turbine section 18.

The gas turbine section 18 as shown in FIG. 2 includes three separate stages. Each stage includes a set of buckets 26, 28, 30, respectively, coupled to respective rotor wheels 32, 34 and 36 that are attached to the turbine rotor or shaft (not shown) in a conventional manner. Between the axially-spaced rows of buckets are annular arrays of stationary blades or vanes 38, 40, 42, respectively, that comprise the turbine nozzles fixed to the surrounding turbine stator (not shown), and labelled S1N, S2N and S3N in FIG. 2.

The rows of nozzles have inner and outer side walls 46, 48, labelled only for the row of vanes 38, but similar inner and outer side walls are associated with the vanes of each nozzle stage. The side walls are typically provided in arcuate-segment form, such that each segment may support one, two or more than two vanes.

FIG. 3 illustrates a vane segment in accordance with a first exemplary but nonlimiting embodiment of the invention. The two vanes 50, 52 have peripheral edge walls 51, 53, respectively, extending about the vanes, connecting the respective leading edges 54, 56 and trailing edges (one shown at 55). The vanes are supported between the inner and outer side wall 46, 48 of the segment. The leading edges 54 56 of the nozzle vanes 50, 52 are exposed to high temperature combustion gases flowing from the transition pieces 24 into the first turbine stage. Since the vanes 50, 52 are substantially identical, only one need be described in detail, it being understood that all of the vanes 38 in the S1N nozzle may be provided with the cooling enhancement described below.

In one exemplary but nonlimiting embodiment, the leading edge 54 is provided with an additional cooling feature independent of any otherwise conventional internal cooling circuit that may be provided within the vane. With particular reference to FIGS. 4-6, a cooling passage or channel 58 extends radially through the peripheral edge wall 51, along the leading edge 54 of the nozzle vane 50, between a radially outer end 60 and a radially inner end 62 of the vane. In the example shown, the inner end 62 is flared on one side at 64. It will be appreciated that the channel 58 may be drilled or cast in place, and that the channel may extend through the inner and outer sidewalls 46, 48. The channel 58 is open at its radially outer and radially inner ends, permitting compressor discharge or extraction air to flow through the channel. Because the channel 58 in the exemplary embodiment is shown extending within the wall thickness of the peripheral edge wall 51, it may be necessary to reinforce the edge wall to maintain a minimum required thickness, as described in greater detail below. It will be appreciated that there may be one cooling channel 58 as shown, or multiple cooling channels spanning the leading edge area of the vane, and the channels may have varying cross-sectional shapes.

FIGS. 7-9 illustrate examples of multiple cooling channel arrangements with different cross-sectional shapes, as well as internal reinforcement rib configurations. Specifically, FIG. 7 shows an array of three cooling passages or channels, including the channel 58 at the leading-most portion of the leading edge 54, along with adjacent channels 66 and 68, all of which have “racetrack-shaped” cross sections, with internal ribs 70, 72 and 74 generally aligned with and extending along the respective channels. In operation, secondary cooling flows from the compressor will pass through the radially-oriented passages or channels 58, 66 and 68, reducing the leading edge temperature, which, in turn, results in improved LCF, Creep, and Oxidation capabilities. The ribs 70, 72 and 74 serve to reinforce the wall thickness of the leading edge 54 directly opposite the respective cooling channels. FIG. 8 shows a similar arrangement, but where the cooling channels 158, 166 and 168 have round cross-sectional shapes, while the reinforcing ribs 170, 172 and 174 remain substantially as shown in FIG. 7. FIG. 9 shows a similar arrangement but where the cooling channels 258, 266 and 268 have substantially rectangular cross-sectional shapes. Here again, the internal reinforcing ribs 258, 266 and 268 are similar to those shown in FIGS. 7 and 8. It will be appreciated that the cooling channel shape and the rib shape may vary as required to produce the desired cooling. For example, the ribs may taper or otherwise have non-uniform thicknesses along their respective lengths.

Using FIG. 7 as an example, it should also be noted that the existing leading edge cooling cavity 76, which may be part of an otherwise conventional internal vane cooling circuit, is not altered by the presence of leading edge cooling passages 58 66 and 68, except that the ribs 70, 72 and 74 within the cooling cavity 76 may serve to increase the surface area within the cavity, and thus may also enhance cooling.

It will also be appreciated that the number and location of the cooling channels provided in the vane peripheral edge wall may vary. For example, as shown in FIG. 10, and depending on cooling requirements, multiple channels 76 may be formed at spaced locations about all or part of the peripheral edge wall 78, with internal reinforcement ribs provided as required to meet minimum thickness requirements for the peripheral edge wall.

As shown in FIG. 11, individual channels may vary in diameter along their respective lengths. Specifically, the cooling channel 80 formed in the peripheral edge wall 82 has a first diameter portion 84 adjacent the outer side wall 86 and a smaller diameter portion 88 at a location radially between the inner and outer side walls. The transition point between diameters can occur as needed, depending on cooling requirements.

FIG. 12 shows another cooling channel arrangement where the radial lengths of cooling channels vary on either side of the leading edge 90. Specifically, two additional channels 92 and 94 having progressively shorter radial lengths are shown to one side of the leading edge channel 96, it being understood that similar channels may be formed to the other side of the leading edge 90, in a symmetrical or asymmetrical manner. Thus, the channels may terminate at any location between the radially inner and/or outer side walls (for example, from 50% to 100% of the radial length of the vane), and may have one or more outlets per channel connecting to the internal cooling cavity within the vane. In FIG. 12, the leading edge channel 96 is shown to have multiple outlets 98, and this feature may be applied to any of the cooling channel arrangements described herein. Note that the outlets 98 may or may not extend through the reinforcing ribs (not shown in FIG. 12).

In addition, the direction of secondary compressor discharge or extraction flow may be in a radial outward or radial inward direction, and thus would determine the inlet and outlet locations for the channels. In other words, the inlets to the channels may be in (or adjacent) one of the inner and outer side walls.

Note that although the present invention may be described primarily in reference to the first stage of an exemplary land-based gas turbine engine, the invention may be applied to any turbine stage, and, a person of ordinary skill in the art, will also appreciate that embodiments of the present invention also may be used in other turbines, including those used in aircraft, and other types of rotary engines.

While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.

Claims

1. A turbine nozzle vane segment comprising:

one or more nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a peripheral edge wall including leading and trailing edges; and at least one substantially radially-oriented cooling channel formed in said peripheral edge wall with an inlet provided at one of said inner and outer side walls.

2. The turbine nozzle vane segment of claim 1 wherein said at least one substantially radially-oriented cooling channel comprises plural, radially-oriented cooling channels including one cooling channel at a forward-most portion of said leading edge.

3. The turbine nozzle vane segment of claim 1 wherein said nozzle vane is substantially hollow, and wherein a radially-oriented rib is provided on an inside surface of said peripheral edge wall, adjacent and extending along said at least one substantially radially-oriented cooling channel.

4. The turbine nozzle vane segment of claim 2 wherein said nozzle vane is substantially hollow, and wherein a radially-oriented rib is provided on an inside surface of said peripheral edge wall, adjacent extending along each of said plural, substantially radially-oriented cooling channels.

5. The turbine nozzle vane segment of claim 4 wherein said radially-oriented rib is located within an internal cooling cavity in said nozzle vane.

6. The turbine nozzle vane segment of claim 1 wherein said at least one radially-oriented cooling channel comprises multiple cooling channels at spaced locations about substantially all of said peripheral edge wall between said leading and trailing edges and wherein a radially-oriented rib is provided on an inside surface of said peripheral edge wall, adjacent and extending along each of said multiple cooling channels.

7. The turbine nozzle vane segment of claim 2 wherein said plural radially-oriented cooling channels have outlets at different radial locations between said inner and outer side walls.

8. The turbine nozzle vane segment of claim 1 wherein each said vane comprises an internal cooling circuit independent of, and not connected to, said at least one substantially radially-oriented cooling channel, and an outlet to said at least one channel is located in the other of said inner and outer side walls.

9. The turbine nozzle vane segment of claim 1 wherein said at least one substantially radially-oriented cooling channel has a round, rectangular or racetrack cross-sectional shape.

10. The turbine nozzle vane segment of claim 1 where at least one substantially radially-oriented cooling channel is formed with different diameter portions along a radial length dimension.

11. A turbine engine comprising a compressor, at least one combustor and at least one turbine stage including a row of stationary nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a peripheral edge wall including leading and trailing edges; and at least one substantially radially-oriented cooling channel formed in said peripheral edge wall with an inlet provided at one of said inner and outer side walls.

12. The turbine engine of claim 10 wherein said at least one substantially radially-oriented cooling channel comprises plural, radially-oriented cooling channels including one cooling channel at a forward-most portion of said leading edge.

13. The turbine engine of claim 11 wherein said nozzle vane is substantially hollow, and wherein a radially-oriented rib is provided on an inside surface of said peripheral edge wall, adjacent and extending along said at least one substantially radially-oriented cooling channel.

14. The turbine engine of claim 12 wherein said nozzle vane is substantially hollow, and wherein a radially-oriented rib is provided on an inside surface of said peripheral edge wall, adjacent and extending along each of said plural, substantially radially-oriented cooling channels.

15. The turbine engine of claim 11 wherein said radially-oriented rib is located within an internal cooling cavity in said nozzle vane.

16. The turbine engine of claim 11 wherein said at least one radially-oriented cooling channel comprises multiple cooling channels at spaced locations about substantially all of said peripheral edge wall between said leading and trailing edges and wherein a radially-oriented rib is provided on an inside surface of said peripheral edge wall, adjacent and extending along each of said multiple cooling channels.

17. The turbine engine of claim 15 wherein said at least one radially-oriented cooling channel has an outlet connected to said internal cooling cavity.

18. The turbine engine of claim 11 wherein said vane comprises an internal cooling circuit independent of, and not connected to, said at least one substantially radially-oriented cooling channel and an outlet to said at least one channel is located in the other of said inner and outer side walls.

19. The turbine engine of claim 11 wherein said at least one substantially radially-oriented cooling channel has a round, rectangular or racetrack cross-sectional shape.

20. A turbine engine comprising a compressor, at least one combustor and at least one turbine stage including a row of stationary nozzle vanes extending between radially inner and outer side walls, each nozzle vane having a peripheral edge wall including leading and trailing edges; and a plurality of substantially radially-oriented cooling channels formed about said peripheral edge wall including at said leading edge, said plurality of substantially radially-oriented cooling channels extending at least part way along a radial length between said inner and outer side walls and having inlets in or adjacent one of said inner and outer side walls; and a reinforcing rib extending along an inside surface of said peripheral edge wall adjacent each of said plurality of substantially radially-oriented cooling channels.