COOLING CONFIGURATION FOR A TURBINE BLADE INCLUDING A SERIES OF SERPENTINE COOLING PATHS

TECHNICAL FIELD

This invention is directed generally to turbine blades and, more particularly, to a turbine blade having a cooling fluid configuration including a series of connected serpentine cooling passages for cooling the pressure side, suction side, and inner walls of the airfoil.

BACKGROUND ART

A conventional gas turbine engine includes a compressor, a combustor, and a turbine. The compressor compresses ambient air, which is supplied to the combustor where the compressed air is combined with a fuel and ignited to create combustion products defining a working gas. The working gas is supplied to the turbine where the gas passes through a plurality of paired rows of stationary vanes and rotating blades. The rotating blades are coupled to a shaft and disc assembly. As the working gas expands through the turbine, the working gas causes the blades, and therefore the shaft and disc assembly, to rotate.

Combustors often operate at high temperatures that may exceed 2500°

Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling fluid configurations for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.

Typically, turbine blades comprise a root, a platform, and an airfoil that extends radially outwardly from the platform. Cooling fluid configurations within most blades include one or more cooling fluid channels that provide cooling to the internal and/or external walls of the airfoil. Such cooling fluid channels may receive air from the compressor of the engine. SUMMARY OF INVENTION

In accordance with a first aspect of the invention, a turbine blade is provided comprising an airfoil including an outer wall extending radially outwardly from a platform to a blade tip. The outer wall includes a pressure sidewall and a suction sidewall, the pressure and suction sidewalls being joined together at chordally spaced apart leading and trailing edges of the airfoil. A pressure side serpentine cooling path includes a plurality of radially extending pressure side cooling channels adjacent the pressure sidewall, wherein the pressure side serpentine cooling path receives cooling fluid for cooling the airfoil from a cooling fluid feed chamber located at least partially radially inwardly from the platform. A suction side serpentine cooling path includes a plurality of radially extending suction side cooling channels adjacent the suction sidewall, wherein the suction side serpentine cooling path receives cooling fluid from the pressure side serpentine cooling path. An intermediate serpentine cooling path includes a plurality of radially extending intermediate cooling channels located between the pressure side cooling channels and the suction side cooling channels, wherein the intermediate serpentine cooling path receives cooling fluid from the suction side serpentine cooling path.

At least a majority of the cooling fluid in the suction side serpentine cooling path may be received from the pressure side serpentine cooling path after the cooling fluid passes through at least a majority of the pressure side cooling channels. For example, substantially all of the cooling fluid in the suction side serpentine cooling path may be received from the pressure side serpentine cooling path after the cooling fluid passes through each of the pressure side cooling channels. At least a majority of the cooling fluid in the intermediate serpentine cooling path may be received from the suction side serpentine cooling path after the cooling fluid passes through at least a majority of the suction side cooling channels.

The turbine blade may further comprise a leading edge cooling system comprising at least one radially extending leading edge pressure side cooling channel adjacent the pressure sidewall of the airfoil near the leading edge, and at least one radially extending leading edge suction side cooling channel adjacent the suction sidewall of the airfoil near the leading edge. The leading edge pressure and suction side cooling channels may each receive cooling fluid from the cooling fluid feed chamber. A spent cooling fluid chamber may extend radially within the airfoil between the leading edge pressure and suction side cooling channels. The spent cooling fluid chamber may comprise a plurality of radially spaced apart cooling fluid outlets located at the leading edge of the airfoil for discharging cooling fluid out of the spent cooling fluid chamber.

At least some of the pressure and suction side cooling channels may include heat conducting pin fins extending from the respective pressure or suction sidewall to a respective inner wall defining the corresponding pressure or suction side cooling channel.

Cooling fluid in the pressure side serpentine cooling path may flow toward the trailing edge of the airfoil, cooling fluid in the suction side serpentine cooling path may flow toward the leading edge of the airfoil, and cooling fluid in the intermediate serpentine cooling path may flow toward the trailing edge of the airfoil.

The turbine blade may further comprise a cooling fluid collection chamber located at least partially radially inwardly from the platform. The cooling fluid collection chamber receives cooling fluid from the pressure side serpentine cooling path and provides at least a portion of the cooling fluid to the suction side serpentine cooling path.

A final one of the intermediate cooling channels may include a plurality of radially spaced apart cooling fluid outlet passages located at or near the trailing edge of the airfoil for discharging cooling fluid out of the intermediate serpentine cooling path. The turbine blade may further comprise at least one cooling fluid exit passage located in another of the intermediate cooling channels and extending to the suction sidewall of the airfoil for discharging cooling fluid from the intermediate serpentine cooling path to the suction sidewall.

The pressure side serpentine path may include at least four pressure side cooling channels provided in series with one another. The suction side serpentine path may include at least four suction side cooling channels in series with one another and being provided downstream from the pressure side cooling channels with reference to a flow of cooling fluid through the airfoil. The intermediate serpentine path may include at least five intermediate cooling channels in series with one another and being provided downstream from the suction side cooling channels with reference to the flow of cooling fluid through the airfoil, thus effecting at least thirteen cooling channels provided in series with one another in the airfoil.

In accordance with a second aspect of the invention, a turbine blade is provided comprising an airfoil including an outer wall extending radially outwardly from a platform to a blade tip. The outer wall includes a pressure sidewall and a suction sidewall, the pressure and suction sidewalls being joined together at chordally spaced apart leading and trailing edges of the airfoil. A pressure side serpentine cooling path includes a plurality of radially extending pressure side cooling channels adjacent the pressure sidewall, wherein the pressure side serpentine cooling path receives cooling fluid for cooling the airfoil from a cooling fluid feed chamber located at least partially radially inwardly from the platform. A suction side serpentine cooling path includes a plurality of radially extending suction side cooling channels adjacent the suction sidewall, wherein the suction side serpentine cooling path receives cooling fluid from the pressure side serpentine cooling path. At least a majority of the cooling fluid in the suction side serpentine cooling path is received from the pressure side serpentine cooling path after the cooling fluid passes through at least a majority of the pressure side cooling channels. An intermediate serpentine cooling path includes a plurality of radially extending intermediate cooling channels located between the pressure side cooling channels and the suction side cooling channels, wherein the intermediate serpentine cooling path receives cooling fluid from the suction side serpentine cooling path. At least a majority of the cooling fluid in the intermediate serpentine cooling path is received from the suction side serpentine cooling path after the cooling fluid passes through at least a majority of the suction side cooling channels. Cooling fluid in the pressure side serpentine cooling path flows toward the trailing edge of the airfoil, cooling fluid in the suction side serpentine cooling path flows toward the leading edge of the airfoil, and cooling fluid in the intermediate serpentine cooling path flows toward the trailing edge of the airfoil. The turbine blade may further comprise a leading edge cooling system comprising at least one radially extending leading edge pressure side cooling channel adjacent the pressure sidewall of the airfoil near the leading edge, and at least one radially extending leading edge suction side cooling channel adjacent the suction sidewall of the airfoil near the leading edge. The leading edge pressure and suction side cooling channels may each receive cooling fluid from the cooling fluid feed chamber. A spent cooling fluid chamber may extend radially within the airfoil between the leading edge pressure and suction side cooling channels. The spent cooling fluid chamber may comprise a plurality of radially spaced apart cooling fluid outlets located at the leading edge of the airfoil for discharging cooling fluid out of the spent cooling fluid chamber.

The turbine blade may further comprise a cooling fluid collection chamber located at least partially radially inwardly from the platform. The cooling fluid collection chamber receives cooling fluid from the pressure side serpentine cooling path and provides at least a portion of the cooling fluid to the suction side serpentine cooling path.

A final one of the intermediate cooling channels may include a plurality of radially spaced apart cooling fluid outlet passages located at or near the trailing edge of the airfoil for discharging cooling fluid out of the intermediate serpentine cooling path. The turbine blade may further comprise at least one cooling fluid exit passage located in another of the intermediate cooling channels and extending to the suction sidewall of the airfoil for discharging cooling fluid from the intermediate serpentine cooling path to the suction sidewall.

The pressure side serpentine path may include at least two pressure side cooling channels provided in series with one another. The suction side serpentine path may include at least two suction side cooling channels in series with one another and being provided downstream from the pressure side cooling channels with reference to a flow of cooling fluid through the airfoil. The intermediate serpentine path may include at least three intermediate cooling channels in series with one another and being provided downstream from the suction side cooling channels with reference to the flow of cooling fluid through the airfoil, thus effecting at least seven cooling channels provided in series with one another in the airfoil.

At least some of the pressure and suction side cooling channels may include: heat conducting pin fins extending from the respective pressure or suction sidewall to a respective inner wall defining the corresponding pressure or suction side cooling channel; and turbulator ribs for increasing turbulence in the flow of cooling fluid passing through the respective pressure and suction side cooling channels.

BRIEF DESCRIPTION OF DRAWINGS

While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:

Fig. 1 is a perspective view of a turbine blade incorporating the present invention;

Fig. 2 is a cross-sectional view of the turbine blade shown in Fig. 1 taken along line 2-2 from Fig. 1 ;

Fig. 3 is a cross-sectional view of the turbine blade shown in Fig. 2 taken along the curved line 3-3 from Fig. 2;

Fig. 4 is a cross-sectional view of the turbine blade shown in Fig. 2 taken along the curved line 4-4 from Fig. 2;

Fig. 5 is a cross-sectional view of the turbine blade shown in Fig. 2 taken along the curved line 5-5 from Fig. 2.

DESCRIPTION OF EMBODIMENTS

In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

Referring to Fig. 1 , an exemplary turbine blade 10 for a gas turbine engine is illustrated. The blade 10 includes an airfoil 12 and a root 14, which is used to secure the blade 10 to a rotor disk (not shown) of the engine for supporting the blade 10 in a hot gas flow path F _P of the engine where hot combustion gases exert motive forces on the surfaces of the blade 10 to provide rotation of the blade 10 and corresponding rotor disk. The airfoil 12 comprises an outer wall 16 including a generally concave pressure sidewall 18 and a generally convex suction sidewall 20. The pressure and suction sidewalls 18, 20 are joined together along a leading edge 22 and a trailing edge 24 of the airfoil 12. The leading and trailing edges 22, 24 are spaced apart from one another in a chordal direction C _D . The airfoil 12 extends outwardly from a platform 26 in a spanwise or radial direction R _D to a blade tip 28 adjacent a radially outer end portion 12B of the airfoil 12, the platform 26 being located adjacent a radially inner end portion 12A of the airfoil 12 as shown in Fig. 1 .

Referring now to Figs. 2 and 3, the airfoil 12 comprises a cooling

configuration 30 including a pressure side serpentine cooling path 36 comprising a plurality of pressure side cooling channels 38a, 38b, 38c, and 38d. The pressure side cooling channels 38a-d are located adjacent the pressure sidewall 18 and extend spanwise (radially) between the radially inner and radially outer end portions 12A, 12B of the airfoil 12. The pressure side cooling channels 38a-d are defined between the pressure sidewall 18, which defines an outer wall of the pressure side cooling channels 38a-d, and a pressure side chordal partition 40 extending chordally through the airfoil 12, which pressure side chordal partition 40 defines an inner wall of the pressure side cooling channels 38a-d. In the illustrated

embodiment, the first channel 38a is separated from the second channel 38b by a first pressure side partition 42, the second channel 38b is separated from the third channel 38c by a second pressure side partition 44, and the third channel 38c is separated from the fourth channel 38d by a third pressure side partition 46. It is noted that the pressure side serpentine cooling path 36 could include additional or fewer pressure side cooling channels than the four channels illustrated in Figs. 2 and 3.

Referring now to Figs. 2 and 4, the cooling configuration 30 further includes a suction side serpentine cooling path 50 comprising a plurality of suction side cooling channels 52a, 52b, 52c, and 52d. The suction side cooling channels 52a-d are located adjacent the suction sidewall 20 and extend spanwise (radially) between the radially inner and radially outer end portions 12A, 12B of the airfoil 12. The suction side cooling channels 52a-d are defined between the suction sidewall 20, which defines an outer wall of the suction side cooling channels 52a-d, and a suction side chordal partition 54 extending chordally through the airfoil 12, which suction side chordal partition 54 defines an inner wall of the suction side cooling channels 52a-d. In the illustrated embodiment, the first channel 52a is separated from the second channel 52b by a first suction side partition 56, the second channel 52b is separated from the third channel 52c by a second suction side partition 58, and the third channel 52c is separated from the fourth channel 52d by a third suction side partition 60. It is noted that the suction side serpentine cooling path 50 could include additional or fewer suction side cooling channels than the four channels illustrated in Figs. 2 and 4.

The pressure side serpentine cooling path 36 and the suction side serpentine cooling path 50 each include a plurality of pin fins 61 , also known in the art as pedestals, as shown in Figs. 2-4. The pin fins 61 in the pressure side serpentine cooling path 36 extend from the interior surface of the pressure sidewall 18 to the pressure side chordal partition 40, and the pin fins 61 in the suction side serpentine cooling path 50 extend from the interior surface of the suction sidewall 20 to the suction side chordal partition 54. The pin fins 61 remove heat from the outer wall 16 of the airfoil 12, which outer wall 16 is heated by the hot combustion gases in the hot gas flow path F _P during operation, and transfer respective portions of the heat from the outer wall 1 6 to each of the pressure and suction side chordal partitions 40, 54. The pin fins 61 also increase turbulence in cooling fluid, such as compressor discharge air, passing through the pressure and suction side serpentine paths 36, 50, thus increasing heat transfer from the outer wall 16 to the pressure and suction side chordal partitions 40, 54. It is noted that the configuration, e.g., number, orientation, distribution, width, etc., of pin fins 61 in each of the cooling channels 38a-d, 52a-d could vary from the configuration shown in Figs. 2-4.

In accordance with an aspect of the present invention, the pressure and suction sidewalls 18, 20 and the pressure and suction side chordal partitions 40, 54 are all of a relatively thin-wall construction, which thin walls are spaced apart from one another by the respective pin fins 61 .

With reference now to Figs. 2 and 5, the cooling configuration 30 further includes an intermediate serpentine cooling path 64 comprising a plurality of intermediate cooling channels 66a, 66b, 66c, 66d, and 66e. The intermediate cooling channels 66a-e are located between the pressure side cooling channels 38a-d and the suction side cooling channels 52a-d, i.e., between the pressure and suction side chordal partitions 40, 54, and extend spanwise (chordally) between the radially inner and radially outer end portions 12A, 12B of the airfoil 12. In the illustrated embodiment, the first channel 66a is separated from the second channel 66b by a first intermediate partition 67, the second channel 66b is separated from the third channel 66c by a second intermediate partition 68, the third channel 66c is separated from the fourth channel 66d by a third intermediate partition 69, and the fourth channel 66d is separated from the fifth channel 66e by a fourth intermediate partition 70. It is noted that the intermediate serpentine cooling path 64 could include additional or fewer intermediate cooling channels than the five channels illustrated in Figs. 2 and 5. As shown in Fig. 5, the fifth channel 66e includes a plurality of pin fins 61 , which extend between the interior surfaces of the pressure and suction sidewalls 18, 20 to define the fifth channel 66e therebetween.

As shown in Figs. 2 and 5, each of the intermediate cooling channels 66a-e may include a plurality of turbulator ribs 71 , also known in the art as trip strips, for increasing turbulence in the flow of cooling fluid passing through the intermediate cooling channels 66a-e. The pressure and suction side cooling channels 38a-d and 52a-d could include similar turbulator ribs 71 as shown in Figs. 3 and 4 to increase turbulence in the flows of the cooling fluid passing through the pressure and suction side cooling channels 38a-d and 52a-d. It is noted that the configuration, e.g., number, orientation, distribution, angle, length, width, etc., of turbulator ribs 71 in each of the cooling channels 38a-d, 52a-d, and 66a-e could vary from the configuration shown in Figs. 2-5.

The intermediate serpentine cooling path 64 may optionally include one or more intermediate cooling fluid outlets 81 extending from at least one of the intermediate cooling channels 66a-e to the suction sidewall 20 of the airfoil 12, see Figs. 2 and 5. In the illustrated embodiment, the first intermediate cooling channel 66a includes a plurality of radially spaced apart intermediate cooling fluid outlets extending to the suction sidewall 20, although other configurations for the cooling fluid outlets 81 could be used. The intermediate cooling fluid outlets 81 are provided to bleed off a portion of the cooling fluid passing through the intermediate serpentine cooling path 64, which cooling air bled off from the intermediate serpentine cooling path 64 via the intermediate cooling fluid outlet 81 may provide film cooling to the suction sidewall 20.

The intermediate serpentine cooling path 64 further comprises a plurality of radially spaced apart trailing edge outlet passages 90 extending from the fifth intermediate cooling channel 66e (the last one of the intermediate cooling channels in the illustrated embodiment) to the trailing edge 24 of the airfoil 12. The trailing edge outlet passages 90 discharge the cooling fluid from the intermediate serpentine cooling path 64 to the trailing edge 24 of the airfoil 12, wherein the cooling fluid provides convective cooling to the trailing edge 24 while passing through the trailing edge outlet passages 90.

As shown in Fig. 2, the cooling configuration 30 further comprises a leading edge cooling system 72 including at least one leading edge pressure side cooling channel 74 (one leading edge pressure side cooling channel 74 is included in the illustrated embodiment), at least one leading edge suction side cooling channel 76 (one leading edge suction side cooling channel 76 is included in the illustrated embodiment), and at least one leading edge spent cooling fluid chamber 78 (one leading edge spent cooling fluid chamber 78 is included in the illustrated embodiment). The leading edge pressure side cooling channel 74 is located adjacent the pressure sidewall 1 8, near the leading edge 22 of the airfoil 12, and extends spanwise (radially) between the radially inner and radially outer end portions 12A, 12B of the airfoil 12. The leading edge suction side cooling channel 76 is located adjacent the suction sidewall 20, near the leading edge 22 of the airfoil 12, and extends spanwise (radially) between the radially inner and radially outer end portions 12A, 12B of the airfoil 12. The leading edge spent cooling fluid chamber 78 is located between the leading edge pressure side cooling channel 74 and the leading edge suction side cooling channel 76, i.e. , between the pressure and suction side chordal partitions 40, 54, and extends spanwise (radially) between the radially inner and radially outer end portions 12A, 12B of the airfoil 12.

The leading edge pressure side cooling channel 74 and the leading edge suction side cooling channel 76 each include a plurality of pin fins 61 as shown in Figs. 2-4. The pin fins 61 in the leading edge pressure side cooling channel 74 extend from the interior surface of the pressure sidewall 18 to the pressure side chordal partition 40, and the pin fins 61 in the leading edge suction side cooling channel 76 extend from the interior surface of the suction sidewall 20 to the suction side chordal partition 54. The pin fins 61 remove heat from the outer wall 16 of the airfoil 12, which outer wall 16 is heated by the hot combustion gases in the hot gas flow path F _P during operation as noted above, and transfer respective portions of the heat to each of the pressure and suction side chordal partitions 40, 54. The pin fins 61 also increase turbulence in cooling fluid passing through the leading edge pressure and suction side cooling channels 74, 76, thus increasing heat transfer from the outer wall 1 6 to the pressure and suction side chordal partitions 40, 54. It is noted that the configuration, e.g. , number, orientation, distribution, width, etc., of pin fins 61 in each of the leading edge pressure and suction side cooling channels 74, 76 could vary from the configuration shown in Figs. 2-4.

The leading edge spent cooling fluid chamber 78 includes a plurality of turbulator ribs 71 for increasing turbulence in the flow of cooling fluid passing through the leading edge spent cooling fluid chamber 78. The leading edge pressure and suction side cooling channels 74, 76 could also include such turbulator ribs 71 as shown in Fig. 3 and 4.

The leading edge spent cooling fluid chamber 78 further comprises a plurality of radially spaced apart cooling fluid outlets 80 extending from the leading edge spent cooling fluid chamber 78 to the leading edge 22 of the airfoil 12. The cooling fluid outlets 80 discharge cooling fluid from the leading edge spent cooling fluid chamber 78 to the leading edge 22 of the airfoil 12, wherein the cooling fluid provides convective cooling to the leading edge 22 while passing through the cooling fluid outlets 80 and provides film cooling for the airfoil 12, e.g., the leading edge 22 and downstream pressure and suction sidewalls 18, 20, upon exiting the cooling fluid outlets 80.

As shown in Fig. 2, a radial dividing wall 82 separates the first intermediate cooling channel 66a of the intermediate serpentine cooling path 64 from the leading edge spent cooling fluid chamber 78 of the leading edge cooling system 72.

Use of the cooling configuration 30 to cool the airfoil 12 during operation of the engine will now be described with reference to Figs. 1 -5.

A cooling fluid, such as compressor discharge air as noted above, may be provided to the airfoil 12 via a cooling fluid feed chamber 86 (see Fig. 5) located at least partially radially inwardly the platform 26. The cooling fluid feed chamber 86 may comprise a single cavity or multiple cavities, which may communicate with one another or may be isolated from one another, located in or around the root 14 of the blade 10. Cooling fluid flows from the cooling fluid feed chamber 86 into the first pressure side cooling channel 38a of the pressure side serpentine cooling path 36 through a pressure side inlet opening 87A, see Fig. 3. The cooling fluid passes sequentially in alternating spanwise directions through the first, second, third, and fourth pressure side cooling channels 38a-d while flowing in a downstream chordal direction as shown in Fig. 3, which downstream chordal direction is defined from the leading edge 22 to the trailing edge 24 of the airfoil 12. The cooling fluid provides cooling to the walls defining the pressure side cooling channels 38a-d and to the blade tip 28 while flowing through the pressure side cooling channels 38a-d. The cooling fluid passes out of the pressure side serpentine cooling path 36 through a pressure side outlet opening 87B and into a cooling fluid collection chamber 88 located at least partially radially inwardly the platform 26, see Fig. 5.

The cooling fluid flows from the cooling fluid collection chamber 88 though a suction side inlet opening 92A (see Fig. 5) into the first suction side cooling channel 52a of the suction side serpentine cooling path 50. The cooling fluid passes sequentially in alternating spanwise directions through the first, second, third, and fourth suction side cooling channels 52a-d while flowing in an upstream chordal direction, i.e., in an opposite chordal direction as the cooling fluid passing through the pressure side serpentine cooling path 36. The cooling fluid provides cooling to the walls defining the suction side cooling channels 52a-d and to the blade tip 28 while flowing through the suction side cooling channels 52a-d. The cooling fluid passes out of the suction side serpentine cooling path 50 through a suction side outlet opening 92B (see Fig. 5) and into the first intermediate serpentine cooling channel 66a of the intermediate serpentine cooling path 64.

A portion of the cooling fluid exits the first intermediate serpentine cooling channel 66a through the intermediate cooling fluid outlets 81 to provide film cooling to the suction sidewall 20 as noted above. The remaining cooling fluid in the intermediate serpentine cooling path 64, which comprises a majority of the cooling fluid in the intermediate serpentine cooling path 64 as only a small portion of the cooling fluid exits the airfoil 12 through the intermediate cooling fluid outlets 81 , passes sequentially in alternating spanwise directions through the first, second, third, fourth, and fifth intermediate cooling channels 66a-e while flowing in a downstream chordal direction, i.e., in the same chordal direction as the cooling fluid passing through the pressure side serpentine cooling path 36. The cooling fluid flows into the last intermediate cooling channel 66e and is discharged from the airfoil 12 through the trailing edge outlet passages 90 where the cooling fluid provides cooling to the trailing edge 24 of the airfoil 12 as discussed above.

In accordance with an aspect of the present invention, since the cooling fluid passing through the cooling configuration 30 is hotter when it enters the

intermediate serpentine cooling path 64 as compared to when the cooling fluid enters the pressure and then the suction side serpentine cooling paths 36, 50, i.e., since the cooling fluid is effectively heated as it passes through the pressure and suction side serpentine cooling paths 36, 50, the temperature of the cooling fluid passing into the intermediate serpentine cooling path 64 may be hotter than the walls defining the intermediate cooling channels 66a-e. Hence, the cooling fluid passing through the intermediate cooling channels 66a-e may effectively heat the walls defining the intermediate cooling channels 66a-e as opposed to cooling them. By heating the walls defining the intermediate cooling channels 66a-e, these walls may become a closer thermal match to the remaining walls of the airfoil 12, e.g., the outer wall 1 6, so as to reduce the thermal gradient between the inner and outer walls of the airfoil 12. Hence, thermal expansion/contraction of these walls may be reduced, thus resulting in reduced stresses on the airfoil inner and outer walls so as to increase the useful lifespan of the blade 10.

With reference to Figs. 2-5, additional portions of the cooling fluid in the cooling fluid feed chamber 86 flow into the leading edge pressure and suction side cooling channels 74, 76 through respective pressure and suction side inlet apertures 94A, 94B (see Figs. 3 and 5). These cooling fluid portions each provide convective cooling as they travel radially outwardly through the leading edge pressure and suction side cooling channels 74, 76 and further provide impingement cooling to the blade tip 28 near the leading edge 22 as they strike against the blade tip 28.

After impinging on the blade tip 28, these portions of cooling fluid flow through respective pressure and suction side outlet apertures 96A, 96B (see Figs. 3 and 5) into the leading edge spent cooling fluid chamber 78. The combined cooling fluid portions then travel radially inwardly through the leading edge spent cooling fluid chamber 78 and exit the leading edge spent cooling fluid chamber 78 through the leading edge cooling fluid outlets 80, wherein the cooling fluid provides cooling to the leading edge 22 of the airfoil 12 as discussed above.

While the number of channels shown in the illustrated embodiment includes four pressure side cooling channels 38a-d, four suction side cooling channels 52a- d, and five intermediate cooling channels 66a-e, for a total of thirteen cooling channels in series with one another, any different number of channels could be used in any of the pressure side serpentine cooling path 36, the suction side serpentine cooling path 50, and the intermediate serpentine cooling path 64, e.g., any combination of channels could be used. For example, an embodiment including a pressure side serpentine cooling path 36 with two pressure side cooling channels, a suction side serpentine cooling path 50 with two suction side cooling channels, and an intermediate serpentine cooling path 64 with three intermediate cooling channels for a total of seven cooling channels in series with one another is contemplated.

While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.