DE102014114244A1 - Gas turbine blade with improved cooling - Google Patents

Abstract

A turbine vane segment includes one or more vanes extending between a radially inner and outer sidewall, each vane having a peripheral edge wall extending between a leading edge and a trailing edge of the vane. In an 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 around the peripheral edge wall, and the inner void space of the vane may be ribbed along the one or more cooling channels and adjacent thereto to reinforce the wall and also around additional cooling surfaces to create the inner cavity.

Description

BACKGROUND TO THE INVENTION

The present application generally relates to improving the efficiency and / or operation of turbines. In particular, but not limited to, the present application relates to improving the cooling of turbine vanes or blades.

A gas turbine 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 blades axially stacked in stages. Each stage may include alternating rows of circumferentially spaced stator vanes that are stationary and circumferentially spaced rotor blades mounted on a wheel fixedly connected to the turbine rotor. In operation, the blades in the compressor rotate with the rotor to compress an airflow supplied to the compressor. Most of the compressed air in the one or more combustion chambers is mixed with a gaseous or liquid fuel and ignited to provide a stream of hot gases which is expanded through the turbine section of the turbine and causes rotation of the turbine rotor. Thus, the energy contained in the fuel is converted into mechanical energy of the rotating rotor, which can be used to rotate the compressor blades to realize the compressed air supply required for combustion, as well as to rotate the coils of a generator that electric power is generated.

Because of the extreme temperatures in the hot gas path, the speed of the working fluid, and the rotational speed of the turbine during operation, the rotating blades (or airfoils) and the stationary vanes are heavily loaded due to extreme mechanical and thermal stresses.

As one skilled in the art will appreciate, there is a strategy for reducing the thermal stresses in the cooling of the vanes and / or the blades so that the temperatures experienced by the vanes and / or blades are lower than those of the hot gas path. Effective cooling can e.g. allow the hot gas path components to withstand higher firing temperatures, withstand higher thermo-mechanical loads and / or extend service life at higher operating temperatures, all of which may allow the turbine to be more cost effective and effective. One way to cool the vanes and blades during operation is to use internal cooling channels or circuits. This generally involves passing relatively cool air that can be supplied from the compressor through internal cooling circuits within the vanes or airfoils.

However, there remains a need for more effective and efficient cooling, e.g. with respect to the leading edges of the stationary vanes exposed to hot combustion gases, especially in the first turbine stage, where the highest temperatures and greatest thermal stresses are experienced.

BRIEF DESCRIPTION OF THE INVENTION

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

The at least one substantially radially oriented cooling passage of the turbine vane segment may include a plurality of radially aligned cooling channels including a cooling passage at a foremost portion of the leading edge.

The vane of any turbine vane segment described above may be substantially hollow, and a radially directed rib may be provided on an inner surface of the peripheral edge wall disposed adjacent to and extending along the at least one substantially radially oriented cooling channel.

The vane of any of the aforementioned turbine nozzle segments may be substantially hollow, with a radially directed rib provided on and extending along an inner surface of the peripheral edge wall adjacent each of the plurality of substantially radially oriented cooling channels.

The radially directed rib of any of the vanes described above may be disposed within an inner cooling cavity in the vane.

The at least one radially directed cooling channel of any one mentioned above Turbine vane segment may include a plurality of cooling channels disposed at spaced locations substantially all around the peripheral edge wall between the leading and trailing edges, wherein a radially aligned rib may be provided on and extending along an inner surface of the peripheral edge wall adjacent each of the plurality of cooling channels ,

The plurality of radially aligned cooling channels of any of the aforementioned turbine nozzle segments may have outlets at various radial locations between the inner and outer sidewalls.

Each vane of any of the aforementioned turbine nozzle segments may have an internal cooling circuit independent of and connected to the at least one substantially radially directed cooling channel, and an outlet of the at least one channel is disposed in the other of the inner and outer side walls.

The at least one substantially radially oriented cooling channel of any of the aforementioned turbine nozzle segment may have a round, rectangular, or racetrack shaped cross-sectional shape.

At least one substantially radially oriented cooling passage of any of the aforementioned turbine nozzle segment may be provided with different diameter sections along a radial length dimension.

In a further non-limiting aspect, the invention provides a turbine including a compressor, at least one combustor and at least one turbine stage having a series of stationary vanes extending between a radially inner and outer wall, each vane having a front and rear peripheral wall Trailing edge, and having at least one substantially radially oriented cooling channel formed in the peripheral edge wall, wherein an inlet is provided on either the inner or on the outer side wall.

The at least one substantially radially oriented cooling passage of the aforementioned turbine may include a plurality of radially directed cooling passages including a cooling passage at a forwardmost portion of the leading edge.

The vane of any of the aforementioned turbines may be substantially hollow, with a radially directed rib on an inner surface of the peripheral edge wall adjacent to and extending along the at least one substantially radially oriented cooling channel.

The vane of any of the aforementioned turbines may be substantially hollow, and a radially directed rib may be provided on and extend along an inner surface of the peripheral edge wall adjacent each of the plurality of substantially radially aligned cooling channels.

The radially directed rib of any of the aforementioned turbines may be disposed within an internal cooling cavity in the vane.

The at least one radially directed cooling channel of any of the aforementioned turbines may include a plurality of cooling channels at spaced locations around substantially the entire peripheral edge wall between the leading and trailing edges, wherein a radially aligned rib on an inner surface of the peripheral edge wall adjacent each of the plurality of radially aligned cooling channels be provided and can extend along this.

The at least one radially directed cooling passage of any of the aforementioned turbines may include an outlet connected to the inner cooling cavity.

The vane of any of the aforementioned turbines may have an internal cooling circuit independent of and connected to the at least one substantially radially directed cooling channel, and an outlet of the at least one channel is in the other of the inner and outer side walls.

The at least one substantially radially oriented cooling channel of any of the aforementioned turbines may have a round, rectangular, or racetrack shaped cross-sectional shape.

It is also a feature of the invention to provide reinforcing ribs in the inner cavity of the vane, adjacent the channels in the peripheral edge wall and at least partially therealong. Accordingly, in yet another exemplary aspect, there is provided a turbine including a compressor, at least one combustor, and at least one turbine stage having a series of stationary vanes extending between a radially inner and outer sidewall, each vane having a front and rear peripheral wall Trailing edge and a plurality of substantially radially aligned cooling channels, which are formed around the peripheral edge wall, including at the front edge, wherein the a plurality of substantially radially aligned cooling channels extend at least a portion along a radial length between the inner and outer side wall and in or adjacent to one of the inner and outer side wall inlets, and having a reinforcing rib, which is adjacent to an inner surface of the peripheral edge wall extending each of the plurality of substantially radially aligned cooling channels.

Another aspect of the invention is a turbine comprising a compressor, at least one combustor, and at least one turbine stage having a series of stationary vanes extending between a radially inner and outer sidewall, each vanes having a front and rear peripheral edge wall and a plurality of substantially radially aligned cooling channels formed around the peripheral edge wall including at the leading edge, wherein the plurality of substantially radially aligned cooling channels extend at least a portion along a radial length between the inner and outer side walls and inlets in or adjacent one of inner and outer side walls and having a reinforcing rib extending along an inner surface of the peripheral edge wall adjacent to each of the plurality of substantially radially aligned cooling channels.

BRIEF DESCRIPTION OF THE DRAWINGS

1 shows a schematic diagram of a conventional gas turbine;

2 shows a partial side view of a turbine section of a conventional gas turbine;

3 FIG. 12 is a perspective view of a gas turbine nozzle set according to an exemplary but not limiting embodiment of the invention; FIG.

4 shows a partially cutaway view illustrating a leading edge cooling channel according to the exemplary, but non-limiting embodiment of the invention;

5 and 6 show enlarged details that out 4 are taken;

7 shows a partial perspective view illustrating the exemplary cooling channel according to an embodiment;

8th shows a partial perspective view illustrating the exemplary cooling channel according to another embodiment;

9 shows a partial perspective view illustrating the exemplary cooling channel according to a still further embodiment;

10 shows a fragmentary plan view of a vane, which is provided with cooling channels according to another exemplary, but not limiting embodiment;

11 shows a partial sectional view of a vane having a cooling channel with different diameters according to another exemplary embodiment; and

12 shows a partial sectional view of a vane showing an array of cooling channels adjacent a leading edge of the vane according to another exemplary embodiment.

DETAILED DESCRIPTION OF THE INVENTION

1 illustrates in schematic form a gas turbine system 10 which, as described hereinafter, comprises stationary vanes and rotating vanes or blades which may be provided with internal cooling circuits. In this otherwise conventional arrangement is through the inlet 12 supplied air in a compressor 14 pressurized and with a fuel in one or more combustion chambers 16 where it is ignited to thereby produce hot combustion gases. Energy becomes the combustion gases in one or more turbine stages located downstream of the combustor 18 taken to a generator 20 to power, which generates electricity. The extracted energy can also be used to drive the compressor 14 used, and note that the turbine rotor 22 the compressor, the turbine stages and the generator can be common. However, the invention described herein is not limited only to the illustrated gas turbine system. Further, in this regard, the cooling circuits described herein are fully compatible with a variety of film cooling configurations that utilize air flowing through cooling circuit passages or cavities.

With additional reference to 2 In a typical gas turbine configuration, multiple combustors may be used 16 annular around the turbine rotor with several transition pieces 24 be arranged, which are the hot combustion gases from the respective combustion chambers to the gas turbine section 18 conduct.

The gas turbine section 18 as he is in 2 is shown has three separate stages. Each stage has a set of blades 26 . 28 respectively. 30 on that with the respective wheels 32 . 34 and 36 coupled to the turbine rotor (not shown) or the turbine shaft are mounted in a conventional manner. Between the axially spaced rows of blades are annular arrays of stationary vanes or vanes 38 . 40 respectively. 42 having the turbine nozzles, which are fixedly connected to the surrounding turbine stator (not shown) and in 2 labeled S1N, S2N and S3N.

The rows of vanes have inner and outer sidewalls 46 . 48 on that only for the series of vanes 38 but similar inner and outer sidewalls belong to the vanes of each vane stage. The sidewalls are typically created in an arcuate segmental shape so that each segment can support one, two, or more than two vanes.

3 FIG. 12 illustrates a vane segment according to a first exemplary but non-limiting embodiment of the invention. FIG. The two vanes 50 . 52 have peripheral edge walls 51 respectively. 53 extending around the vanes and the respective leading edges 54 . 56 and trailing edges (one of which at 55 shown). The blades are between the inner and the outer wall 46 . 48 of the segment held. The leading edges 54 . 56 the vanes 50 . 52 are exposed to high temperature combustion gases from the transition pieces 24 flow into the first turbine stage. Because the vanes 50 . 52 are essentially identical, only one must be described in detail, it being understood that all the vanes 38 in the nozzle S1N may be provided with the below-described cooling improvement.

In an exemplary but non-limiting embodiment, the leading edge is 54 provided with an additional cooling device, which is independent of any otherwise conventional internal cooling circuits and which may be provided within the guide vane. With particular reference to FIGS 4 - 6 extends a cooling passage or channel 58 radially through the peripheral edge wall 51 through, along the leading edge 54 the vane 50 between a radially outer end 60 and a radially inner end 62 the vane. In the example shown, the inner end 62 at one side 64 widened. It is recognized that the channel 58 can be drilled or cast in place and that the channel extends through the inner and outer sidewall 46 . 48 can extend. The channel 58 is open at its radially outer and radially inner ends and allows the compressor to flow exhaust or branch air through the duct. Because the channel 58 in the exemplary embodiment, as being within the wall thickness of the peripheral edge wall 51 extending, it may be necessary to reinforce the edge wall to maintain a minimum required thickness, as described in more detail below. It will be appreciated that, as shown, a single cooling channel 58 or more cooling channels may span the leading edge region of the vane, and the channels may have varying cross-sectional shapes.

7 - 9 illustrate examples of multiple cooling channel assemblies having different cross-sectional shapes as well as inner reinforcing rib configurations. In particular shows 7 an array of three cooling passages or channels, including the channel 58 at the foremost portion of the leading edge 54 , together with adjacent channels 66 and 68 all of which have a "racetrack" section with internal ribs 70 . 72 and 74 which are substantially aligned with and extend along the respective channels. During operation, secondary cooling flows from the compressor pass through the radially aligned passageways or channels 58 . 66 and 68 which reduces the temperature of the leading edge, which in turn leads to improved low frequency fatigue strength, creep resistance and oxidation resistance. Ribs 70 . 72 and 74 serve to the wall thickness of the leading edge 54 directly opposite the respective cooling channels to reinforce. 8th shows a similar arrangement in which the cooling channels 158 . 166 and 168 but have round cross-sectional shapes, while the reinforcing ribs 170 . 172 and 174 essentially as in 7 stay shown. 9 shows a similar arrangement, but in which the cooling channels 258 . 266 and 268 have substantially rectangular cross-sectional shapes. Here are the inner reinforcing ribs again 270 . 272 and 274 those in the 7 and 8th shown similar. It will be appreciated that the cooling channel shape and rib shape may vary as required to provide the desired cooling. For example, the ribs may be tapered or otherwise have a non-uniform thickness along their respective length.

When using the 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 inner vane cooling circuit, by the presence of the leading edge passageways 58 . 66 and 68 not changed, except that the ribs 70 . 72 and 74 inside the cooling cavity 76 can serve to increase the surface area within the cavity and thus improve the cooling.

It will further be appreciated that the number and location of the cooling channels provided in the peripheral edge wall of the vane may vary. As in 10 shown and depending on the cooling requirements, for example, multiple channels 76 at spaced locations around all or part of the peripheral edge wall 78 be formed, wherein inner reinforcing ribs, as required, are provided to meet minimum thickness requirements of the peripheral edge wall.

As in 11 As shown, the diameters of individual channels may vary along their respective lengths. In particular, the cooling channel 80 which is in the peripheral edge wall 82 is formed, a first diameter section 84 adjacent to the outer sidewall 86 and a section 88 with a smaller diameter at a location radially between the inner and outer sidewalls. The transition point between the diameters may occur as required depending on cooling requirements.

12 shows a further cooling channel arrangement in which the radial lengths of the cooling channels on each side of the leading edge 90 vary. In particular, on one side of the leading edge channel 96 two additional channels 92 and 94 shown having progressively shorter radial lengths, it being understood that similar channels are in a symmetrical or asymmetrical manner on the other side of the leading edge 90 can be formed. Thus, the channels may terminate at any location between the radially inner and / or outer sidewalls (eg, from 50% to 100% of the radial length of the vane) and have one or more outlets per channel that communicate with the inner cooling cavity within the vane are connected. In 12 is the leading edge channel 96 with several outlets 98 and this feature can be applied to any cooling channel arrangement described herein. Note that the outlets 98 through the (in 12 but not shown) can extend through reinforcing ribs.

Additionally, the direction of the secondary output or branch flow of the compressor may be directed radially outward or radially inward, and thus would determine the location of the inlets and outlets for the channels. In other words, the inlets to the channels may be located in (or adjacent to) either the inner or outer sidewall.

Note that while the present invention is described primarily with respect to the first stage of an exemplary land-based gas turbine, the invention can be applied to any turbine stage, and one skilled in the art will also recognize that embodiments of the present invention are also applicable to other turbines, including those used in de aviation and other types of rotary machines can be used.

While the invention is described in conjunction 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 changes and equivalent arrangements. which are included within the scope and scope of the appended claims.

A turbine vane segment includes one or more vanes extending between a radially inner and outer sidewall, each vane having a peripheral edge wall extending between a leading edge and a trailing edge of the vane. In an 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 around the peripheral edge wall, and the inner void space of the vane may be ribbed along the one or more cooling channels and adjacent thereto to reinforce the wall and also around additional cooling surfaces to create the inner cavity.

LIST OF REFERENCE NUMBERS

10

 turbine system

12

 inlet

14

 compressor

16

 combustors

18

 turbine stages

20

 generator

22

 turbine rotor

24

 Transitions

26, 28, 30

 blades

 32, 34 and 36

 impellers

38, 40, 42

Shovels or vanes

46, 48, 86

 side walls

50, 52

 vanes

51, 53, 78, 82

 edge walls

54, 56, 90

 leading edge

55

 trailing edge

58, 66, 68, 80, 92, 94, 96, 158, 166, 168, 258, 266 and 268

Cooling passages or channels

60

outer end

62

inner end

64

(expanded) page

70, 72 and 74

Inner ribs

170, 172, 174, 270, 272 and 274

 reinforcing ribs

76

 cooling cavity

84, 88

 Diameter portions

90

 leading edge

98

 outlets

Claims (10)

Turbine vane segment comprising: one or more vanes extending between a radially inner and outer side wall, each vane having a peripheral edge wall having leading and trailing edges; and at least one substantially radially directed cooling passage formed in the peripheral edge wall with an inlet provided at one of the inner and outer side walls. 2. The turbine nozzle vane segment of claim 1, wherein the at least one substantially radially oriented cooling channel includes a plurality of radially aligned cooling channels including a cooling channel at a forwardmost portion of the leading edge. The turbine nozzle vane segment of claim 1, wherein the vane is substantially hollow and wherein on an inner surface of the peripheral edge wall is provided a radially directed rib adjacent to and extending along the at least one substantially radially oriented cooling channel. 2. The turbine nozzle vane segment of claim 2, wherein the vane is substantially hollow and wherein on an inner surface of the peripheral edge wall is provided a radially directed rib adjacent to and extending along each of the plurality of substantially radially aligned cooling channels. 2. The turbine nozzle vane segment of claim 1, wherein the at least one radially aligned cooling channel has a plurality of cooling channels at spaced locations about substantially the entire peripheral edge wall between the leading and trailing edges and wherein a radially directed rib is provided on an inner surface of the peripheral edge wall and adjacent each the plurality of radially aligned cooling channels and extends along this. 2. The turbine nozzle vane segment of claim 2, wherein the plurality of radially aligned cooling channels have outlets at different radial locations between the inner and outer walls. 2. The turbine nozzle vane segment of claim 1, wherein each vane has an inner cooling circuit independent of and unconnected to the at least one substantially radially directed cooling channel, and wherein an outlet of the at least one channel is disposed in the other of the inner and outer side walls , Turbine vane segment according to claim 1, wherein at least one substantially radially oriented cooling channel is formed along a radial length dimension with different diameter portions. A turbine having a compressor, at least one combustor, and at least one turbine stage having a series of stationary vanes extending between a radially inner and outer sidewall, each vane having a peripheral edge wall having leading and trailing edges and at least one substantially radially oriented cooling channel formed in the peripheral edge wall having an inlet provided at one of the inner or outer side wall. A turbine including a compressor, at least one combustor, and at least one turbine stage having a series of stationary vanes extending between a radially inner and outer sidewall, each vanes having a front and rear edge peripheral wall and a plurality of substantially radially oriented cooling ducts are formed around the peripheral edge wall, including at the leading edge, wherein the plurality of substantially radially aligned cooling channels extend at least a portion along a radial length between the inner and outer sidewalls and have inlets in or adjacent one of the inner and outer sidewalls and a reinforcing rib extending along an inner surface of the peripheral edge wall adjacent to each of the plurality of substantially radially aligned cooling channels.

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