TURBINE BLADE WITH IMPROVED TRAILING EDGE COOLING

Abstract

A turbine blade (110) is provided, where an exterior surface of the turbine blade (110) is exposed to a hot combustion gas (125) above a flow path line (141). The turbine blade (110) includes a trailing edge pin bank cooling channel (118) in an airfoil section (112), a cooling air supply channel (120) in a root section (114), and a channel (122) shaped with a radius of curvature (126), to extend the channel (122) across the flow path line (141) and interconnect the cooling air supply channel (120) to the trailing edge pin bank cooling channel (118). A width of the trailing edge pin bank cooling channel (118) is adjusted, such that the width at an inner diameter region (128) and an outer diameter region (131) is less than the width at an intermediate region (130) between the inner and outer diameter regions (128,131).

Description

FIELD OF THE INVENTION

The present invention relates to turbine blades, and more specifically, to a turbine blade having an improved cooling system.

BACKGROUND OF THE INVENTION

Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade for producing power. Combustors often operate at high temperatures that may exceed 2500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of the exposure to the high temperatures.

FIG. 1 illustrates a conventional turbine blade 10 including an airfoil section 12 joined to a root section 14 at a platform section 16. Generally, the airfoil section 12 includes a pressure sidewall and a suction sidewall (not shown) joined along a leading edge 11 and a trailing edge 13 and extending radially outward from the platform section 16 at an inner diameter region 28 to a tip 21 at an outer diameter region 31. The airfoil section 12 includes a cooling system 19 with a serpentine cooling passage that receives cooling fluid from an inlet in the root section 14. A cooling fluid flow 39 from the root section 14 is received in the serpentine cooling passage, which includes a first channel 44 and a second channel 46 joined by a first turn 45 at the tip 21, and a third channel 48 joined to the second channel 46 by a second turn 47 at the inner diameter region 28. During the gas turbine operation, a hot combustion fluid flow 25 passes around an exterior of the airfoil section 12 above the platform section 16 which defines a flow path line 41. The flow path line 41 defines the radial extent of the area of the blade which is exposed to the hot combustion gas. Additionally, a cooling fluid flow 23 passes from a supply channel 20 into a third channel 48, through parallel impingement orifices 51 and a channel 22 and then into an impingement trailing edge cooling channel 18 along the trailing edge 13, after which the cooling fluid passes through impingement orifices 52 between spaced-apart ribs 58, before exiting the airfoil section 12 through the trailing edge 13.

FIG. 2 is a closer view of the channel 22 that connects the third channel 48 in the airfoil section 12 to the impingement trailing edge cooling channel 18 in the airfoil section 12. As illustrated in FIG. 2, the channel 22 defines a cross-sectional flow area for the cooling fluid at an entrance to the impingement trailing edge cooling channel 18, between a rib 38 and a surface 27 of the airfoil section 12. The surface 27 has a radius of curvature 26. As further illustrated in FIG. 2, the channel 22 is positioned above the flow path line 41 and above the platform section 16 of the turbine blade 10 such that the radius of curvature 26 is located entirely in the region of the blade exposed to the hot combustion gas. In an exemplary embodiment, the channel 22 is a collection chamber through which the cooling fluid 23 passes at a slow or static rate.

FIG. 3 illustrates a conventional ceramic core 64 used to form the conventional turbine blade 10 of FIG. 2 during a casting process, as appreciated by one of skill in the art. The core 64 includes a rib opening 70 corresponding to the rib 38 of the turbine blade 10 and a channel portion 68 with a radius of curvature 26 that corresponds to the channel 22 of the turbine blade 10 subsequent to the casting process. The design of the core 64, including the radial position of the rib opening 70 as well as the radius of curvature 26 of the channel portion 68, are determinative of the cross-sectional flow area of the cooling fluid which will pass through the entrance to the impingement trailing edge cooling channel 18 of the turbine blade 10. The core 64 also includes a channel portion 72 corresponding to the impingement trailing edge cooling channel 18 of the turbine blade 10.

FIG. 8 illustrates the impingement trailing edge cooling channel 18 of the conventional turbine blade 10. Reflecting the tapered width of the blade along its radial axis, the width 32 of the cooling channel 18 (measured between the pressure side 15 and the suction side 17 perpendicular to a radial axis of the blade) at the inner diameter region 28 is greater than the width 34 of the cooling channel at the intermediate region 30, and the width 34 of the cooling channel 18 at the intermediate region 30 is greater than the width 36 of the cooling channel 18 at the outer diameter region 31. Thus, the width 32,34,36 of the cooling channel 18 continuously increases from the inner diameter region 28 to the outer diameter region 31. The turbine blade 10 has an outer thickness 29 at the inner diameter region 28. As appreciated by one of skill in the art, the width of the channel 22, and thus the flow of cooling fluid through the channel 22 and into the cooling channel 18 is responsive to the width of the cooling channel 18.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in the following description in view of the drawings that show:

FIG. 1 is a cross-sectional side view of a conventional turbine blade;

FIG. 2 is a partial view of the cross-sectional side view of FIG. 1 at an entrance to an impingement trailing edge cooling channel;

FIG. 3 is a cross-sectional side view of a conventional core;

FIG. 4 is a top view of a turbine blade in accordance with an embodiment of the present invention;

FIG. 5 is a cross-sectional side view of the turbine blade of FIG. 4 along the line 5-5;

FIG. 6 is a cross-section side view of a core used to form the turbine blade of FIG. 5;

FIG. 7 is a cross-sectional end view of the turbine blade of FIG. 4 along the line 7-7; and

FIG. 8 is a cross-sectional end view of the turbine blade of FIG. 1 along the line 8-8

DETAILED DESCRIPTION OF THE INVENTION

The present inventors have recognized several limitations of the prior art blade designs. For example, the inventors have recognized that conventional ceramic casting cores 64 are prone to breakage at the radial extremities, such as along the channel portion 68 where sudden changes in stiffness occur. Furthermore, the inventors have recognized that the cooling system 19 does not distribute the cooling fluid across the radial dimension of the cooling channel 18 in a manner consistent with the heat transfer from the combustion fluid flow 25 across the radial dimension of the turbine blade 10. Heat load is typically a maximum proximate to a mid-span region of the airfoil. The ceramic core is prone to breakage in the channel portion 68. To improve the core 64 strength, the channel 22 is enlarged in a manner which causes a disproportionate amount of cooling flow to enter the cooling channel 18 through the channel 22, where the heat load is relatively low, thereby providing a mismatch between the heat load and the cooling capacity.

Thus, the present inventors have developed a blade with an improved cooling arrangement which provides improved cooling performance while at the same time being castable with a ceramic core which is less prone to breakage at the radial extremities. The flow of cooling fluid across the radial dimension of the turbine blade in the present invention is more closely matched to the heat load distribution from the combustion fluid flow across the radial dimension of the turbine blade exterior.

In order to address the shortcomings of the conventional core design, the present inventors recognized that the radial extremities are prone to breakage, due to a low radius of curvature of the surfaces of the core in those regions, which result in sudden changes in stiffness and a concentration of stress. The present inventors recognized that in an improved core design, the radial extremities would be reshaped to increase the radius of curvature of the radial extremity surfaces in order to reduce or eliminate the sudden changes in stiffness in those critical regions.

Additionally, the present inventors recognized that the radius of curvature of the channel portion of the core, which shapes the radius of curvature of the channel of the turbine blade during the casting process, should be increased. With the improved core design, the increased radius of curvature of the channel portion results in an extension of the channel of the turbine blade from the airfoil section, across the flow path line, through the platform section and into the root section.

By increasing the radius of curvature of the channel of the turbine blade, a cross-sectional flow area through the entrance to the trailing edge impingement cooling channel increases. To maintain a desired total cross-sectional area, in addition to increasing the radius of curvature of the channel portion, the core design is further modified to move a rib opening toward the root in the radial direction, which corresponds to lowering of a rib of the turbine blade in the radial direction, such that the increase in the cross-sectional flow area in the radial direction caused by an increase in the radius of curvature of the channel is offset to maintain the predetermined flow area. Additionally, to further improve the cooling system of the turbine blade such that the cooling effectiveness of cooling fluid at the inner and outer diameter regions is less than at an intermediate region between the inner and outer diameter regions, the core design is modified such that the width of the trailing edge impingement cooling channel in the turbine blade is greater at the intermediate region than at the inner and outer diameter regions. The modified width profile of the trailing edge impingement cooling channel more closely matches the heat exchange capacity of the cooling fluid to the heat load imposed across the radial extent of the blade.

FIGS. 4, 5 and 7 illustrate a turbine blade 110 in accordance with one embodiment of the invention, including a leading edge 111, a trailing edge 113, a pressure side 115 and a suction side 117. FIG. 5 illustrates that the turbine blade 110 includes an airfoil section 112, a root section 114, and a platform section 116 positioned between the airfoil section 112 and the root section 114. The turbine blade 110 includes a cooling system 119 positioned between the pressure side 115 and the suction side 117, and includes several channels to pass a cooling fluid which is delivered from the root section 114. The cooling system 119 includes a serpentine network of channels 144,146,148 including a first channel 144 which extends along the leading edge 111 from an inner diameter region 128 to an outer diameter region 131 of the airfoil section 112. At the outer diameter region 131, a first turn (not shown) is provided, to join the first channel 144 to a second channel 146, which extends from the outer diameter region 131 to the inner diameter region 128. At the inner diameter region 128, a second turn 147 joins the second channel 146 with a third channel 148. A hot combustion fluid flow 125 from a combustor (not shown) of a turbine engine passes over the exterior of the airfoil section 112, above a flow path line 141. The flow path line 141 defines the radial extent of the area of the blade 110 which is exposed to the hot combustion fluid flow 125. The temperature distribution of the hot combustion fluid flow 125 in a radial direction over the exterior of the airfoil section 112 is greater in radial regions between the inner diameter region 128 and the outer diameter region 131 than at the inner diameter region 128 and the outer diameter region 131. In an exemplary embodiment, the temperature distribution of the hot combustion gas 125 is a maximum at an intermediate region 130 positioned half-way between the inner diameter region 128 and the outer diameter region 131.

Cooling fluid is provided to the serpentine network of channels 144,146,148 from a supply channel 160 within the root section 114, and passes into the first channel 144 at the inner diameter region 128. The cooling fluid subsequently passes to the first turn (not shown) at the outer diameter region 131, after which the cooling fluid passes into the second channel 146 and then flows from the outer diameter region 131 to the second turn 147 at the inner diameter region 128, after which the cooling fluid passes to the third channel 148. As the cooling fluid passes through the third channel 148, the cooling fluid partially passes through the impingement orifices 150 positioned between segmented ribs 156 aligned between the third channel 148 and the trailing edge pin bank cooling channel 118. The cooling fluid subsequently passes through impingement orifices 152,154 of respectively spaced apart ribs 158,160, before exiting the trailing edge 113 through the orifices 162.

In addition to the serpentine network of channels 144,146,148, cooling fluid may enter the trailing edge pin bank cooling channel 118 by passing from a cooling air supply channel 120 within the root section 114, into a channel 122. The channel 122 begins within the root section 114 and passes across the flow path line 141 before communicating with the trailing edge pin bank cooling channel 118 in the airfoil section 112. The cooling fluid within the trailing edge bin bank cooling channel 118 passes through the orifices 152 of the segmented ribs 158, after which the cooling fluid subsequently passes through the orifices 154 of the segmented ribs 160, and finally passes through the orifices 162 of the trailing edge 113, to exit the airfoil section 112. The turbine blade 110 and the airfoil section 112 illustrated in FIGS. 4-5 and 7 are exemplary, and the turbine blade and airfoil section may have various alternative designs, with varying numbers of channels and/or structural design, and still be within the scope of the embodiments of the present invention.

As illustrated in FIG. 5, the channel 122 interconnects the cooling air supply channel 120 to the trailing edge pin bank cooling channel 118. The channel 122 receives the cooling fluid from the cooling air supply channel 120 within the root section 114 and passes the cooling fluid through an entrance of the trailing edge pin bank cooling channel 118 within the airfoil section 112. As illustrated in FIG. 5, a surface 127 of the turbine blade 110 adjacent to the inner diameter region 128 is shaped with a radius of curvature 126. The channel 122 is shaped, based on the surface 127, and thus the channel 122 is similarly shaped with the radius of curvature 126 that defines a curvature of the channel 122. FIG. 6 illustrates a core 164 used to form the turbine blade 110 during a casting process, as appreciated by one of skill in the art. The radius of curvature 126 of the channel 122 and the surface 127 are defined based on the radius of curvature 126 of a corresponding channel portion 168 of the core 164 during the casting process. Thus, the channel 122 extends from the cooling air supply channel 120 in the root section 114, through the platform section 116, and into the trailing edge pin bank cooling channel 118 in the airfoil section 112, based on a radius of curvature 126 of the channel portion 168 of the core 164. Alternatively, one may describe the cooling channel 118 as extending via a radius of curvature to below a flow path line of the blade. This is counterintuitive to prior art blade designs where it was known to terminate a trailing edge cooling channel at or above the flow path line, since there is no heat load being supplied into the blade material below the flow path line.

FIG. 3 illustrates the conventional core 64 used to form the conventional turbine blade 10 of FIGS. 1-2, including a channel portion 68 with the radius of curvature 26 that is used to shape the channel 22 in the conventional turbine blade 10. As previously discussed, the channel portion 68 of the conventional core 64 is prone to breakage, due to the low radius of curvature 26 which results in sudden changes in stiffness across the channel portion 68 and thus concentrated stress in this region. As illustrated in FIG. 6, the design of the core 164 used to form the turbine blade 110 features notable variations from the design of the conventional core 64 used to form the conventional turbine blade 10. As illustrated in FIGS. 3 and 6, the radius of curvature 126 of the channel portion 168 of the core 164 is greater than the radius of curvature 26 of the channel portion 68 for the conventional core 64 (for an equivalent core design), to reduce or eliminate sudden changes in stiffness across the channel portion 168, and thus locations of concentrated stress. Thus, the core 164 is less prone to breakage across the channel portion 168, as compared to the core 64. Additionally, the increased radius of curvature 126 of the channel portion 168 of the core 164 will increase the radius of curvature 126 of the channel 122 of the turbine blade 110, such that the channel 122 extends from the cooling air supply channel 120 in the root section 114 to the trailing edge pin bank cooling channel 118 in the airfoil section 112. As illustrated in FIGS. 1-2, the channel 22 of the conventional turbine blade 10, shaped by the channel portion 68 of the conventional core 64, extends within the airfoil section 12 from the third channel 48 to the cooling channel 18, and does not extend below the flow path line 41 or extend into the root section 14.

As previously discussed, in order to improve the cooling system 119 of the turbine blade 110, the flow of cooling fluid at the inner and outer diameter regions 128,131 is reduced below the flow of cooling fluid at the intermediate region 130 between the inner and outer diameter regions 128,131. Also, as previously discussed, the core 164 is designed, such that the radius of curvature 126 of the channel 122 is increased, which consequently increases the entrance to the cooling channel 118, in a radial direction, thereby increasing the flow of cooling fluid through the entrance to the cooling channel 118 at the inner diameter region 128. In order to offset this increased flow, the increased entrance to the cooling channel 118 in the radial direction is offset, such that the flow area through the entrance to the cooling channel 118 is mitigated or limited to a predetermined flow area, thereby mitigating a flow of cooling fluid through the entrance to the cooling channel 118 at the inner diameter region 128. To effect this offset, the rib opening 170 of the core 164 is lowered (relative to the rib opening 70 of the conventional core 64), such that the rib 138 of the turbine blade 110 is radially positioned to cooperate with the radius of curvature 126 of the channel 122, to offset the increased entrance to the cooling channel 118 in the radial direction, and maintain the predetermined cross-sectional flow area through the channel 122 at the entrance of the trailing edge pin bank cooling channel 118. By increasing the radius of curvature 126 of the channel portion 168 in the core 164 (from the conventional core 64 design), the curvature of the channel 122 is reduced, to extend the channel 122 from the airfoil section 112 through the platform section 116 and into the root section 114, and subsequently increase the cross-sectional flow area through the channel 122 in the radial direction at the entrance to the cooling channel 118. If the rib 138 is not radially lowered in the airfoil section 112 design, to offset the increase in the cross-sectional flow area in the radial direction, the inner diameter region 128 of the airfoil section 112 would be overcooled by an excessive cooling fluid flow through the channel 122 at the entrance to the cooling channel 118. Thus, by designing the core 164 such that the rib opening 170 is appropriately positioned, the cross-sectional flow area through the channel 122 at the entrance to the cooling channel 118 is maintained at a predetermined flow rate, to enhance a cooling efficiency of the cooling system 119 of the turbine blade 110.

As previously discussed, FIG. 8 illustrates the trailing edge cooling channel 18 of the conventional turbine blade 10, in which the width 32 (measured between the pressure side 15 and the suction side 17) at the inner diameter region 28 is greater than the width 34 at the intermediate diameter region 30 positioned between the inner and outer diameter regions 28,31. Thus, the flow of cooling fluid through the inner diameter region 28 of the conventional turbine blade 10 exceeds the flow of cooling fluid flow through the intermediate diameter region 30. The efficiency of the cooling system 19 arrangement is limited, as the temperature distribution of the hot combustion fluid flow 25 is greater at the intermediate region 30 than at the inner and outer diameter regions 28,31, resulting in an overcooled inner and outer diameter regions 28,31 and an undercooled intermediate region 30. The core 164 used to form the turbine blade 110 is designed, to further improve the cooling system 119 of the turbine blade 110, to avoid these shortcomings of the cooling system 19 of the conventional turbine blade 10. The channel portion 172 of the core 164 is used to form the trailing edge cooling channel 118 of the turbine blade 110. The channel portion 172 of the core 164 is reshaped, to adjust the width of the trailing edge pin bank cooling channel 118 (FIG. 7), which extends in a radial direction from the inner diameter region 128 attached to the platform section 116, to the outer diameter region 131. As further illustrated in FIG. 7, the channel portion 172 of the core 164 is reshaped, such that a width 132,136 of the trailing edge pin bank cooling channel 118 at the inner diameter region 128 and at the outer diameter region 131 is less than the width 134 of the trailing edge pin bank cooling channel 118 at a mid-span region or an intermediate region 130 positioned between the inner diameter region 128 and the outer diameter region 131. In an exemplary embodiment, the intermediate region 130 is positioned at a midpoint or half-way between the inner diameter region 128 and the outer diameter region 131. Although FIG. 7 illustrates that the maximum width is the width 134 of the trailing edge pin bank cooling channel 118 at the intermediate region 130 positioned half-way between the inner and outer diameter regions 128, 131, the embodiments of the present invention is not limited to this arrangement, provided that the width of the trailing edge pin bank cooling channel 118 is greater at a region between the inner and outer diameter regions 128,131 than the width 132,136 at the respective inner and outer diameter regions 128,131. The redesigned core 164 and channel portion 172 are used to cast the cooling channel 118 of the turbine blade 110 which conforms to the distribution of the heat temperature of the combustion fluid flow 125 over the exterior of the airfoil section 112, during the operation of the turbine engine. By adjusting the width 132,136 of the cooling channel 118 at the inner and outer diameter regions 128,131 to be less than the width 134 of the cooling channel 118 at the intermediate region 130, the flow of cooling fluid through the inner and outer diameter regions 128,131 is mitigated, thereby increasing the flow of cooling fluid at the intermediate region 130. Similarly, as previously discussed, the design of the core 164 is configured to shape the entrance to the cooling channel 118 in a radial dimension, to also mitigate the flow of cooling fluid through the inner diameter region 128 of the cooling channel 118. The increased radius of curvature 126 of the channel portion 168 of the core 164 increased the cross-sectional flow area (in the radial dimension) through the entrance to the cooling channel 118. However, the core 164 design also featured lowering of the rib opening 170, such that the increased cross-sectional flow area (in the radial dimension) is offset through the entrance to the cooling channel 118, to mitigate the flow of cooling fluid through the inner diameter region 128.

Thus, the core 164 design enhances the cooling efficiency of the cooling system 119, as the core 164 design is configured to shape the cooling channel 118 such that the radial dimension of the entrance to the cooling channel 118 is mitigated or limited, to mitigate the flow of cooling fluid through the inner diameter region 128 of the cooling channel (and consequently that the flow of cooling fluid through the intermediate region 130 is increased). Similarly, the core 164 design enhances the cooling efficiency of the cooling system 119, as the core 164 design shapes the cooling channel 118, in a lateral dimension, so that the width 132,134 at the inner and outer diameter regions 128,131 are less than the width 136 at the intermediate region 130, to mitigate the flow of cooling fluid at the inner and outer diameter regions 128,131 and enhance the flow of cooling fluid at the intermediate region 130. Thus, the design of the core 164, and the trailing edge pin bank cooling channel 118 enhances the cooling efficiency of the trailing edge pin bank cooling channel 118 and the cooling system 119, by enhancing the flow of cooling fluid at the intermediate region 130 and mitigating the flow of cooling fluid at the inner and outer diameter regions 128,131. Although FIG. 7 illustrates a design to be used with regard to the trailing edge pin bank cooling channel 118, an equivalent channel design may be utilized with regard to the other channels 144,146,148 of the cooling system 119 or in channels used in cooling systems other than the cooling system 119, for example.

Although FIG. 5 illustrates an embodiment of a turbine blade 110 design in which the channel 122 is shaped to extend from the cooling air supply channel 120 in the root section 114 to the trailing edge pin bank cooling channel 118, and FIG. 7 illustrates an embodiment of a trailing edge pin bank cooling channel 118 design, the embodiments of FIGS. 5 and 7 need not be incorporated into the same turbine blade 110. Thus, the turbine blade may be designed with the channel 122 design of FIG. 5, but without the trailing edge pin bank cooling channel 118 design of FIG. 7. Similarly, the turbine blade may be designed with the trailing edge pin bank cooling channel 118 design of FIG. 7, but without the channel 122 design of FIG. 5.

While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.

Claims

1. A turbine blade including a trailing edge cooling channel wherein the improvement comprises the trailing edge cooling channel having a maximum width proximate a mid-span region of an airfoil section of the blade and extending via a radius of curvature to below a flow path line of the blade.

2. A turbine blade including an airfoil section, a root section, and a platform section between the airfoil section and the root section, wherein an exterior surface of said turbine blade is exposed to a hot combustion gas above a flow path line, said turbine blade comprising:

a trailing edge pin bank cooling channel in the airfoil section;

a cooling air supply channel in the root section;

a channel interconnecting the cooling air supply channel to the trailing edge pin bank cooling channel, said channel having a radius of curvature to extend the channel from the cooling air supply channel in the root section, across the flow path line and to the trailing edge pin bank cooling channel in the airfoil section.

3. The turbine blade of claim 2, wherein said airfoil section extends in a radial direction from an inner diameter region attached to the platform section, to an outer diameter region; and wherein a width of the trailing edge pin bank cooling channel at the inner diameter region and at the outer diameter region is less than a width of the trailing edge pin bank cooling channel at an intermediate region positioned between the inner diameter region and the outer diameter region.

4. The turbine blade of claim 3, wherein said width of the trailing edge pin bank cooling channel is a maximum width at the intermediate region positioned half-way between the inner diameter region and the outer diameter region.

5. The turbine blade of claim 2, further comprising a rib being radially positioned to cooperate with the radius of curvature of the channel to define a predetermined cross-sectional flow area through the channel at an entrance of the trailing edge pin bank cooling channel.

6. A ceramic core used to cast the turbine blade of claim 2 and including a portion defining the channel.

7. A turbine blade including an airfoil section, a root section and a platform section between the airfoil section and the root section, said turbine blade comprising:

a trailing edge pin bank cooling channel in the airfoil section;

wherein said airfoil section extends in a radial direction from an inner diameter region attached to the platform section, to an outer diameter region;

and wherein a width of the trailing edge pin bank cooling channel at the inner diameter region and at the outer diameter region is less than the width of the trailing edge pin bank cooling channel at an intermediate region positioned between the inner diameter region and the outer diameter region.

8. The turbine blade of claim 7, wherein an exterior surface of said turbine blade is exposed to a hot combustion gas above a flow path line, said turbine blade further comprising:

a cooling air supply channel in the root section; and

a channel to interconnect the cooling air supply channel to the trailing edge pin bank cooling channel, said channel having a radius of curvature to extend the channel from the cooling air supply channel in the root section across the flow path line and to the trailing edge pin bank cooling channel in the airfoil section.

9. The turbine blade of claim 7, wherein said width of the trailing edge pin bank cooling channel is a maximum width at the intermediate region positioned half-way between the inner diameter region and the outer diameter region.

10. The turbine blade of claim 8, further comprising a rib being radially positioned to cooperate with the radius of curvature of the channel to define a predetermined cross-sectional flow area through the channel at an entrance of the trailing edge pin bank cooling channel.

11. A ceramic core used to cast the turbine blade of claim 7.