TURBINE BLADE WITH IMPINGEMENT CAVITY COOLING INCLUDING PIN FINS

Abstract

A turbine blade including an airfoil having a cavity defined between pressure and suction side walls, and first and second cooling circuits located within the cavity. The second cooling circuit includes a first upstream impingement cavity, a second upstream impingement cavity and a discharge impingement cavity. A plurality of pin fins are located within each of the impingement cavities, extending between the pressure and suction side walls to provide cooling to the pressure and suction side walls, and to restrict flow through the second cooling circuit and to direct an increased flow of cooling fluid into the first cooling circuit.

Description

FIELD OF THE INVENTION

This invention is directed generally to turbine blades and, more particularly, to a turbine blade having cooling circuits for conducting a cooling fluid through an airfoil of the blade to provide an improved thermal balance between parts of the airfoil.

BACKGROUND OF THE INVENTION

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 ignites the mixture, creating 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 rotor assembly. As the working gas expands through the turbine, the working gas causes the blades, and therefore the rotor assembly, to rotate.

Typical turbine combustor configurations expose turbine blade assemblies to 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 systems 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 outwardly from the platform. The airfoil typically comprises a tip, a leading edge and a trailing edge. Most blades typically contain internal cooling channels forming a cooling system. The cooling channels in the blades may receive air from the compressor of the turbine engine and pass the air through the blade. The cooling channels often include multiple flow paths that are designed to maintain the turbine blade at a relatively uniform temperature. However, some configurations of flow circuits comprising the flow channels may prevent some areas of the turbine blade from being adequately cooled, which may result in the formation of localized hot spots.

SUMMARY OF THE INVENTION

In accordance with an aspect of the invention, a turbine blade is provided comprising an airfoil including an airfoil outer wall extending in a span-wise direction radially outwardly from a blade root. The airfoil outer wall may further include pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil, wherein the airfoil defines an airfoil cavity forming a cooling system in the blade. A plurality of trailing edge cooling slots are located in a trailing edge wall along the trailing edge and extend from the airfoil cavity to an exterior surface of the outer wall. At least two ribs extend in the span-wise direction and are chordally spaced from each other within the airfoil cavity, each rib defining an upstream side and a downstream side and having a plurality of impingement orifices extending between respective upstream and downstream sides. An upstream impingement cavity is defined between an upstream one of the ribs and a downstream one of the ribs and extends substantially the span of the airfoil, wherein the impingement orifices in the upstream rib provide a flow of accelerated impingement air onto the downstream one of the ribs. A discharge impingement cavity is defined between the downstream one of the ribs and the trailing edge wall and extends substantially the span of the airfoil, wherein the impingement orifices in the downstream one of the ribs provide a flow of accelerated impingement air onto the trailing edge wall. A plurality of pin fins extend between the pressure and suction side walls within the discharge impingement cavity, and a plurality of pin fins extend between the pressure and suction side walls within the upstream impingement cavity.

In accordance with further aspects of the invention, the turbine blade may further comprise a blade tip wall located at an end of the airfoil distal from the root, an outer framing area passage defined between a radially outer end of the downstream one of the ribs and the blade tip wall, an inner framing area passage wall located radially inwardly from a radially inner end of the downstream one of the ribs, and an inner framing area passage defined between the radially inner end of the downstream one of the ribs and the inner framing area passage wall. Further, the turbine blade may include one or more pin fins located in or near at least one of the outer and inner framing area passages adjacent to the downstream one of the ribs such that the one or more pin fins provide resistance to a flow of air flowing in the chordal direction through the at least one of the outer and inner framing area passages.

The turbine blade may further include an air supply passage extending from the blade root to the airfoil cavity. The airfoil cavity may include a first cooling circuit extending from the air supply passage and toward the leading edge, and the airfoil cavity may include a second cooling circuit comprising the upstream and discharge impingement cavities. The first cooling circuit may comprise a serpentine path having a first passage extending radially to the blade tip wall and receiving cooling air from the supply passage. A supply of cooling air for the second cooling circuit may be provided by the first passage of the first cooling circuit. The plurality of pin fins within at least the upstream impingement cavity may restrict flow of cooling air through the second cooling circuit to reduce a flow rate of cooling air through the second cooling circuit and increase a flow rate of cooling air through the first cooling circuit.

At least one of the pin fins in one or more of the impingement cavities may be radially aligned with at least one of the outer and inner framing area passages to restrict flow. Further, at least a portion of at least one of the pin fins may be located within the outer framing area passage and radially aligned with an outermost portion of at least one of the ribs. Additionally, at least a portion of at least one of the pin fins is located within the inner framing area passage and radially aligned with an innermost portion of at least one of the ribs.

In accordance with still further aspects of the invention, the turbine blade may include a first rib located upstream from the upstream one of the ribs and a first impingement cavity defined for the second cooling circuit between the first rib and the upstream one of the ribs, wherein the upstream impingement cavity may comprise an intermediate impingement cavity between the first impingement cavity and the discharge impingement cavity.

The turbine blade may additionally include a plurality of impingement orifices through the first rib to provide a flow of accelerated impingement air onto the upstream one of the ribs. Additionally, the turbine blade may include a plurality of pin fins extending between the pressure and suction side walls within the first impingement cavity.

In accordance with another aspect of the invention, a turbine blade is provided comprising an airfoil including an airfoil outer wall extending in a span-wise direction radially outwardly from a blade root. The airfoil outer wall may further include pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil, wherein the airfoil defines an airfoil cavity forming a cooling system in the blade. A plurality of trailing edge cooling slots are located in a trailing edge wall along the trailing edge and extend from the airfoil cavity to an exterior surface of the outer wall. First, second and third ribs extend in the span-wise direction and are chordally spaced from each other within the airfoil cavity, Each rib defines an upstream side and a downstream side and has a plurality of impingement orifices extending between respective upstream and downstream sides. A first impingement cavity is defined between the first and second ribs and extends substantially the span of the airfoil, wherein the impingement orifices in the first rib provide a flow of accelerated impingement air onto the second rib. An intermediate impingement cavity is defined between the second and third ribs and extends substantially the span of the airfoil, wherein the impingement orifices in the second rib provide a flow of accelerated impingement air onto the third rib. A discharge impingement cavity is defined between the third rib and the trailing edge cooling slots and extends substantially the span of the airfoil, wherein the impingement orifices in the third rib provide a flow of accelerated impingement air onto the trailing edge wall. A plurality of pin fins extend between the pressure and suction side walls within the discharge impingement cavity, and a plurality of pin fins extend between the pressure and suction side walls within at least one of the first impingement cavity and the intermediate impingement cavity. In accordance with a further aspect of the invention, a plurality of pin fins may extend between the pressure and suction side walls within each of the first impingement cavity and the intermediate impingement cavity. In accordance with further aspects of the invention, the turbine blade may comprise a blade tip wall located at an end of the airfoil distal from the root and an outer framing area passage may be defined between a radially outer end of the third rib and the blade tip wall. One or more of the pin fins may be located in or near the outer framing area passage such that the one or more pin fins provide resistance to a flow of air flowing in the chordal direction through the outer framing area passage.

The turbine blade may comprise an inner framing area passage wall located radially inwardly from a radially inner end of the third rib. Further, an inner framing area passage may be defined between the radially inner end of the third rib and the inner framing area passage wall. Still further, the turbine blade may include one or more pin fins located in or near the inner framing area passage such that the one or more pin fins provide resistance to a flow of air flowing in the chordal direction through the inner framing area passage.

The turbine blade may comprise outer and inner framing area passages located adjacent to respective radially outer and inner ends of the second rib. One or more of the pin fins may be located in or near at least one of the outer and inner framing area passages adjacent to the second rib such that the one or more pin fins provide resistance to a flow of air flowing in the chordal direction through the at least one of the outer and inner framing area passages.

An air supply passage may extend from the blade root to the airfoil cavity, and the airfoil cavity may include a first cooling circuit extending from the air supply passage and toward the leading edge and a second cooling circuit comprising the first, intermediate and discharge impingement cavities. The first cooling circuit may comprise a serpentine path having a first passage extending radially to a radially outer end of the airfoil outer wall and receiving cooling air from the air supply passage, wherein a supply of cooling air for the second cooling circuit may be provided from the first passage of the first cooling circuit to the first impingement cavity.

BRIEF DESCRIPTION OF THE 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 partial cross section of a turbine blade according to aspects of the present invention;

FIG. 2 is an enlarged partially cut away cross-sectional view of a trailing edge of an airfoil of the turbine blade shown in FIG. 1; and

FIG. 3 is a perspective view of a partial cross section of the airfoil of the turbine blade according to aspects of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

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.

Reference is now made to FIG. 1 which shows a turbine blade 2 according to aspects of the present invention. The turbine blade 2 is one of a plurality of turbine blades that are circumferentially aligned in a gas turbine engine. The turbine blade 2 comprises an airfoil 4 extending radially outwardly from a mounting location such as a rotor disk of a rotor assembly. The airfoil 4 includes an airfoil outer wall 6 extending in a span-wise direction radially outwardly from a blade platform 8 and blade root 10. The airfoil outer wall 6 is exposed to a high temperature working gas of the gas turbine engine. Additionally, the airfoil outer wall 6 comprises both a pressure side wall 12 on an airfoil outer wall side facing the flow of high temperature working gas and an oppositely facing suction side wall 14. The pressure and suction side walls 12, 14 are joined together at chordally spaced apart respective leading and trailing edges 16, 18 of the airfoil 4. The turbine blade 2 further comprises a blade tip wall 20 located at a radially outer end or blade tip 21 of the airfoil 4 distal from the blade root 10.

Referring again to FIG. 1, the turbine blade airfoil 4 is integrally connected to the turbine blade platform 8 and blade root 10 at a radially inner location of the turbine blade 2. The blade platform 8 comprises a surface 30 that extends generally perpendicular to the span of the airfoil 4. A plurality of blade platforms of the circumferential array of turbine blades 2 forms an inner boundary to a hot gas path through a turbine section of a gas turbine engine. Cooling air is provided through a turbine rotor assembly (not shown) to provide cooling air to components of the gas turbine engine including the turbine blade 2.

Referring now to FIG. 3, the cooling air provided through the turbine rotor assembly to cool the turbine blade 2 is supplied to an airfoil cavity 26. In particular, the airfoil 4 of the turbine blade 2 defines an airfoil cavity 26 forming a cooling system in the turbine blade 2. A main air supply passage 32, a middle air supply passage 34 and a leading edge air supply passage 36 provide a flow of cooling air radially outwardly into the airfoil cavity 26 of the turbine blade 2. Each air supply passage 32, 34, 36 extends from the blade root 10 to the airfoil cavity 26 wherein the blade root 10 includes one or more passages in fluid communication with the cooling air provided through the turbine rotor assembly. The main air supply passage 32 is located closest to the trailing edge 18 of the turbine blade 2 and may supply a larger volume of cooling air than either of the two other air supply passages 34, 36 to the turbine blade airfoil cavity 26. The middle air supply passage 34 and the leading edge air supply passage 36 provide additional cooling air to the airfoil cavity 26 at two forward locations of the airfoil 4, with the leading edge air supply passage 36 shown in FIG. 3 as the forward-most passage.

The airfoil cavity 26 includes a first cooling circuit 38 generally extending from the main air supply passage 32 toward the leading edge 16, and a second cooling circuit 40 generally extending from the main air supply passage 32 toward the trailing edge 18. The first cooling circuit 38 comprises a forward serpentine path 48 having a first passage 50 extending radially outwardly from the main air supply passage to the blade tip wall 20. A second passage 52 of the first cooling circuit 38 is in fluid communication with the first passage 50 at a chordal passage 51 and extends radially inwardly at the blade tip wall 20 toward the platform 8. A third passage 54 of the first cooling circuit 38 is in fluid communication with the second passage 52 at a chordal passage 53 at a radially inner end of the second passage 52 near the platform 8, receiving air from the second passage 52 and combining with air flowing radially outwardly from the middle air supply passage 34, as shown by the arrows in FIG. 3. The cooling air from the first cooling circuit serpentine path 48 and the middle air supply passage 34 exit the turbine blade 2 through outlets (not shown) at the blade tip 21.

As seen in FIG. 3, a leading edge air supply cavity 57 extends from the leading edge air supply passage 36 radially outwardly to the blade tip wall 20. The air supplied to the leading edge air supply cavity 57 comprises an air supply for a leading edge impingement cavity 58 that extends radially along the leading edge 16 of the airfoil 4. The leading edge impingement cavity 58 is defined by a leading edge rib 60 which includes a plurality of leading edge impingement orifices 62 for providing impingement cooling air through the leading edge impingement cavity 58 to the leading edge 16.

A supply of cooling air for the second cooling circuit 40 is provided from the first passage 50 of the first cooling circuit 38. As shown in FIGS. 1 and 2, the second cooling circuit 40 comprises three impingement cavity ribs 64, 66, 68 extending in the span-wise direction and chordally spaced from each other within the airfoil cavity 26. The rib 64 includes an upstream side 64a and a downstream side 64b, the rib 66 includes an upstream side 66a and a downstream side 66b, and the rib 68 includes an upstream side 68a and a downstream side 68b. Further, each rib 64, 66, 68 has a respective plurality of impingement orifices 64c, 66c, 68c.

Referring to FIG. 2, the ribs 64, 66, 68 define three impingement cavities 42, 44, 46 forming passages of the second cooling circuit 40 located near the trailing edge 18 of the airfoil 4, i.e., opposite or downstream from the first cooling circuit 36. A first, upstream impingement cavity 42 extends substantially the span of the airfoil 4 and is defined between the rib 64, comprising a first, upstream rib 64, and the rib 66, comprising a second, upstream rib 66. The impingement orifices 64c extend between upstream and downstream sides 64a, 64b of the first rib 64 and provide a flow of accelerated impingement air onto the second rib 66, and additionally impinge along converging portions of the inner surfaces of the pressure and suction side walls 12, 14 defining the first impingement cavity 42. A second, intermediate impingement cavity 44 extends substantially the span of the airfoil 4 and is defined between the second rib 66 and the rib 68, defining a third, downstream rib 68. The impingement orifices 66c extend between upstream and downstream sides 66a, 66b of the second rib 66 and provide a flow of accelerated impingement air onto the third rib 68, and additionally impinge along converging portions of the inner surfaces of the pressure and suction side walls 12, 14 defining the second impingement cavity 44. A third discharge impingement cavity 46 extends substantially the span of the airfoil 4 and is defined between the third rib 68 and a trailing edge wall 23 defining a plurality of trailing edge cooling slots 22. The impingement orifices 68c extend between upstream and downstream sides 68a, 68b of the third rib 68 and provide a flow of accelerated impingement air onto the trailing edge wall 23 immediately upstream from the trailing edge cooling slots 22, and additionally impinge along converging portions of the inner surfaces of the pressure and suction side walls 12, 14 defining the discharge impingement cavity 46. It may be noted that the trailing edge slots 22 are defined between a plurality of radially spaced dividers 28 forming a portion of the trailing edge wall 23.

As described above, either one or both of the first impingement cavity 42 and the intermediate impingement cavity 44 comprise upstream impingement cavities, i.e., upstream of the discharge impingement cavity 46. Further, either one or both of the first rib 64 and the second rib 66 comprise upstream ribs, i.e., upstream of the downstream rib 68.

As noted previously, the second cooling circuit 40 receives cooling air from the first passage 50 of the first cooling circuit 38, and provides a flow of cooling air in an opposite direction from the first cooling circuit 38 toward the trailing edge 18 of the airfoil 4. In particular, cooling air flows successively through each of the impingement cavities 42, 44, 46 and then passes out of the second cooling circuit 40 through the trailing edge cooling slots 22.

The second cooling circuit 40 further includes outer framing area passages 84, 86, 88 defined between the outer ends of the respective impingement cavity ribs 64, 66, 68 and the blade tip wall 20, and inner framing area passages 92, 94, 96 defined between the inner ends of the respective impingement cavity ribs 64, 66, 68 and an inner framing area wall 90. The inner framing area wall 90 is located radially inwardly from the inner ends of the ribs 64, 66, 68, and may be generally radially aligned with the platform 8. The outer and inner framing area passages 84, 86, 88 and 92, 94, 96 are formed by chordally extending end portions of a core (not shown) for supporting impingement cavity forming portions of the core during a casting process to form the airfoil 4. Additional cooling air may pass from the first passage 50 of the first cooling circuit 38 through the outer and inner framing area passages 84, 86, 88 and 92, 94, 96 to the discharge impingement cavity 46.

Referring to FIGS. 1 and 2, each of the impingement cavities 42, 44, 46 includes a plurality of pin fins, such as micro pin fins, extending between the pressure and suction side walls 12, 14. Specifically, the first upstream impingement cavity 42 includes a plurality of pin fins 78 distributed or spaced from each other in the chordal and radial directions between the blade tip wall 20 and the inner framing area wall 90; the second upstream (intermediate) impingement cavity 44 includes a plurality of pin fins 80 distributed or spaced from each other in the chordal and radial directions between the blade tip wall 20 and the inner framing area wall 90; and the discharge impingement cavity 46 includes a plurality of pin fins 82 distributed or spaced from each other in the chordal and radial directions between the blade tip wall 20 and the inner framing area wall 90. The pin fins 78, 80, 82 are distributed through substantially the entire span of their respective impingement cavities 42, 44, 46. The pin fins 78, 80, 82 are arranged within the cavities 42, 44, 46 such that air passing through respective ones of the impingement orifices 64c, 66c, 68c may impinge on one or more of the respective pin fins 78, 80, 82 to effect convective heat transfer from the pin fins 78, 80, 82, and to deflect and disperse the incoming impingement air within the respective impingement cavities 64, 66, 68. It may be noted that the air passing through the impingement orifices 64c, 66c, 68c is accelerated toward the next successive upstream rib side 66a, 68a and the trailing edge wall 23 to facilitate convective cooling within the impingement cavities 64, 66, 68.

In accordance with an aspect of the invention, in the configuration of the first and second cooling circuits 38, 40 described herein, the pin fins 78, 80, 82 operate to restrict the cooling air flow rate through the second cooling circuit 40, while also providing an increase in the effectiveness of heat transfer from the pressure and suction side walls 12, 14 via the pin fins 78, 80, 82. In particular, it may be noted that in the absence of the pin fins 78, 80, 82, the cooling air flow provided from the first passage 50 of the first cooling circuit 38 to the second cooling circuit 40 would be greater than may be effectively utilized for the required convective heat transfer within the impingement cavities 42, 44, 46, with an associated decrease in the necessary cooling air flow provided to the first cooling circuit 38. The pin fins 78, 80, 82 provide a restriction in the overall flow area for the second cooling circuit 40, and thereby increase the portion of cooling air flow that is distributed through the first cooling circuit 38 toward the leading edge 16, with an associated increase in convective heat transfer through the first cooling circuit 38.

In addition, particular ones of the pin fins 78, 80, 82 are provided in the flow paths formed by the outer and inner framing area passages 84, 86, 88 and 92, 94, 96. For example, as seen in FIG. 2, one or more pin fins 78a and 78b in the first upstream impingement cavity 42 may be located at the radially outer and inner locations in or near the outer and inner framing area passages 84, 86 and 92, 94, respectively, to restrict or slow the flow of cooling air past the ends of the ribs 64, 66; one or more pin fins 80a and 80b in the second upstream (intermediate) impingement cavity 44 may be located at the radially outer and inner locations in or near the outer and inner framing area passages 86, 88 and 94, 96, respectively, to restrict or slow the flow of cooling air past the ends of the ribs 66, 68; and one or more pin fins 82a and 82b in the third discharge impingement cavity 46 may be located at the radially outer and inner locations in or near the outer and inner framing area passages 88 and 96, respectively, to restrict or slow the flow of cooling air past the ends of the rib 68. It should be understood that the radially located pin fins 78a, 78b, 80a, 80b, 82a, 82b are at or near locations that are radially aligned with the framing area passages 84, 86, 88, 92, 94, 96, or may be located at least partially extending between an end of a rib 64, 66, 68 and one of the blade tip wall 20 and the inner framing area wall 90. It should be noted that additional features, such as the protruding portion 93 shown on the blade tip wall 20 adjacent to the radially outer portion of the second rib 66, may be provided as a further restriction to the flow through the framing area passages 84, 86, 88, 92, 94, 96.

The placement of the radially located pin fins 78a, 78b, 80a, 80b, 82a, 82b in or near the framing area passages 84, 86, 88, 92, 94, 96 provides a substantial resistance to fluid flow for restricting or slowing the flow of cooling fluid in the chordal direction past the ends of the ribs 64, 66, 68 such that a larger proportion of the cooling fluid flowing through the second cooling circuit 40 may flow through the impingement orifices 64c, 66c, 68c, as compared to flow through the framing area passages 84, 86, 88, 92, 94, 96, to further improve the effective use of cooling fluid in transferring heat from the pressure and suction side walls 12, 14 in the trailing edge portion of the airfoil 4. The effectiveness of the cooling air flowing through the impingement cavities 64, 66, 68 is additionally increased by the plurality of pin fins 78, 80, 82 providing additional heat transfer for cooling the side walls 12, 14, as described above.

Also, as noted above, the resistance to flow provided by the configuration of pin fins in each of the impingement cavities 42, 44, 46 operates to increase the flow rate in the first cooling circuit 38. The increased flow rate in the first cooling circuit 38 allows the forward serpentine path 48 to be formed with a larger cross-sectional passage area while maintaining a required Mach number in the forward flowing cooling fluid for providing effective cooling to the portions of the pressure and suction side walls 12, 14 located along the first cooling circuit 38.

Although not shown in the FIGS. 1-3, the core design for the manufacturing of the airfoil cavity 26 benefits from the arrangement of pin fins of the present invention. The pin fins 78, 80, 82 are included in the core as holes that extend between two opposing core sides that lighten the core near the trailing edge 18. Lightening the core with pin fin holes has the advantage of reducing the risk of core fracture and improving the core yield.

In accordance with an aspect of the invention, the configuration of the pin fins 78, 80, 82 within the impingement cavities 42, 44, 46 may be utilized to obtain a desired flow in the first and second cooling circuits 38, 40. For example, the diameter of the pin fins 78, 80 82, the number of pin fins 78, 80, 82 and/or the placement of the pin fins 78, 80, 82 within the impingement cavities 42, 44, 46 may be selected to increase or decrease the flow area through the second cooling circuit 40 for controlling the balance of cooling air flow from the first passage 50 to the cooling circuits 38, 40.

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.

Claims

1. A turbine blade comprising:

an airfoil including an airfoil outer wall extending in a span-wise direction radially outwardly from a blade root;

said airfoil outer wall including pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of said airfoil, said airfoil defining an airfoil cavity forming a cooling system in said blade;

a plurality of trailing edge cooling slots located in a trailing edge wall along said trailing edge and extending from said airfoil cavity to an exterior surface of said outer wall;

at least two ribs extending in the span-wise direction and chordally spaced from each other within the airfoil cavity, each said rib defining an upstream side and a downstream side and having a plurality of impingement orifices extending between respective upstream and downstream sides;

an upstream impingement cavity defined between an upstream one of said ribs and a downstream one of said ribs and extending substantially the span of said airfoil, and said impingement orifices in said upstream rib providing a flow of accelerated impingement air onto said downstream one of said ribs;

a discharge impingement cavity defined between said downstream one of said ribs and said trailing edge wall and extending substantially the span of said airfoil, and said impingement orifices in said downstream one of said ribs providing a flow of accelerated impingement air onto said trailing edge wall;

a plurality of pin fins extending between said pressure and suction side walls within said discharge impingement cavity; and

a plurality of pin fins extending between said pressure and suction side walls within said upstream impingement cavity.

2. The turbine blade of claim 1, further comprising:

a blade tip wall located at an end of said airfoil distal from said root;

an outer framing area passage defined between a radially outer end of said downstream one of said ribs and said blade tip wall;

an inner framing area passage wall located radially inwardly from a radially inner end of said downstream one of said ribs, and an inner framing area passage defined between said radially inner end of said downstream one of said ribs and said inner framing area passage wall; and

including one or more pin fins located in or near at least one of said outer and inner framing area passages adjacent to said downstream one of said ribs such that said one or more pin fins provide resistance to a flow of air flowing in the chordal direction through said at least one of said outer and inner framing area passages.

3. The turbine blade of claim 2, including an air supply passage extending from said blade root to said airfoil cavity, said airfoil cavity including a first cooling circuit extending from said air supply passage and toward said leading edge, and including a second cooling circuit comprising said upstream and discharge impingement cavities.

4. The turbine blade of claim 3, wherein said first cooling circuit comprises a serpentine path having a first passage extending radially to said blade tip wall and receiving cooling air from said air supply passage, and wherein a supply of cooling air for said second cooling circuit is provided by said first passage of said first cooling circuit.

5. The turbine blade of claim 4, wherein said plurality of pin fins within at least said upstream impingement cavity restrict flow of cooling air through said second cooling circuit to reduce a flow rate of cooling air through said second cooling circuit and increase a flow rate of cooling air through said first cooling circuit.

6. The turbine blade of claim 2, wherein at least one of said pin fins in one or more of said impingement cavities is radially aligned with at least one of said outer and inner framing area passages to restrict flow.

7. The turbine blade of claim 6, wherein at least a portion of at least one of said pin fins is located within said outer framing area passage and radially aligned with an outermost portion of at least one of said ribs.

8. The turbine blade of claim 6, wherein at least a portion of at least one of said pin fins is located within said inner framing area passage and radially aligned with an innermost portion of at least one of said ribs.

9. The turbine blade of claim 1, including a first rib located upstream from said upstream one of said ribs and defining a first impingement cavity for said second cooling circuit between said first rib and said upstream one of said ribs, and said upstream impingement cavity comprising an intermediate impingement cavity between said first impingement cavity and said discharge impingement cavity.

10. The turbine blade of claim 9, including a plurality of impingement orifices through said first rib providing a flow of accelerated impingement air onto said upstream one of said ribs.

11. The turbine blade of claim 10, including a plurality of pin fins extending between said pressure and suction side walls within said first impingement cavity.

12. A turbine blade comprising:

an airfoil including an airfoil outer wall extending in a span-wise direction radially outwardly from a blade root;

said airfoil outer wall including pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of said airfoil, said airfoil defining an airfoil cavity forming a cooling system in said blade;

a plurality of trailing edge cooling slots located in a trailing edge wall along said trailing edge and extending from said airfoil cavity to an exterior surface of said outer wall;

first, second and third ribs extending in the span-wise direction and chordally spaced from each other within the airfoil cavity, each said rib defining an upstream side and a downstream side and having a plurality of impingement orifices extending between respective upstream and downstream sides;

a first impingement cavity defined between said first and second ribs and extending substantially the span of said airfoil, and said impingement orifices in said first rib providing a flow of accelerated impingement air onto said second rib;

an intermediate impingement cavity defined between said second and third ribs and extending substantially the span of said airfoil, and said impingement orifices in said second rib providing a flow of accelerated impingement air onto said third rib;

a discharge impingement cavity defined between said third rib and said trailing edge wall and extending substantially the span of said airfoil, and said impingement orifices in said third rib providing a flow of accelerated impingement air onto said trailing edge wall;

a plurality of pin fins extending between said pressure and suction side walls within said discharge impingement cavity; and

a plurality of pin fins extending between said pressure and suction side walls within at least one of said first impingement cavity and said intermediate impingement cavity.

13. The turbine blade of claim 12, including a plurality of pin fins extending between said pressure and suction side walls within each of said first impingement cavity and said intermediate impingement cavity.

14. The turbine blade of claim 12, further comprising:

a blade tip wall located at an end of said airfoil distal from said root;

including an outer framing area passage defined between a radially outer end of said third rib and said blade tip wall, and

including one or more of said pin fins located in or near said outer framing area passage such that said one or more pin fins provide resistance to a flow of air flowing in the chordal direction through said outer framing area passage.

15. The turbine blade of claim 14, further comprising:

an inner framing area passage wall located radially inwardly from a radially inner end of said third rib;

an inner framing area passage defined between said radially inner end of said third rib and said inner framing area passage wall; and

including one or more pin fins located in or near said inner framing area passage such that said one or more pin fins provide resistance to a flow of air flowing in the chordal direction through said inner framing area passage.

16. The turbine blade of claim 15, further comprising:

outer and inner framing area passages located adjacent to respective radially outer and inner ends of said second rib; and

including one or more pin fins located in or near at least one of said outer and inner framing area passages adjacent to said second rib such that said one or more pin fins provide resistance to a flow of air flowing in the chordal direction through said at least one of said outer and inner framing area passages.

17. The turbine blade of claim 15, wherein at least one of said pin fins in one or more of said impingement cavities is radially aligned with at least one of said outer and inner framing area passages to restrict flow.

18. The turbine blade of claim 15, wherein at least one of said pin fins in one or more of said impingement cavities is radially aligned with at least one of said outer framing area passages to restrict flow in the chordal direction; and

wherein at least one of said pin fins in one or more of said impingement cavities is radially aligned with at least one of said inner framing area passages to restrict flow in the chordal direction.

19. The turbine blade of claim 12, including an air supply passage extending from said blade root to said airfoil cavity, said airfoil cavity including a first cooling circuit extending from said air supply passage and toward said leading edge, and including a second cooling circuit comprising said first, intermediate and discharge impingement cavities.

20. The turbine blade of claim 19, wherein said first cooling circuit comprises a serpentine path having a first passage extending radially to a radially outer end of said airfoil outer wall and receiving cooling air from said air supply passage, and wherein a supply of cooling air for said second cooling circuit is provided from said first passage of said first cooling circuit to said first impingement cavity.