Title:
Turbine airfoil with integral chordal support ribs
Kind Code:
A1


Abstract:
A turbine airfoil including a structural support system formed from chordal ribs within an internal cooling system in the airfoil. The chordal ribs forming the structural support system may be positioned in a cooling fluid supply channel in the hollow airfoil that is positioned proximate to a trailing edge cooling channel. The trailing edge cooling channel may include a plurality of pin fins forming a pin fin bank extending between inner surfaces of the pressure side wall and suction side wall. The chordal ribs may protrude from an inner surface of the outer wall of the hollow airfoil and extend in a general chordwise direction from a mid-chord region of the hollow airfoil to the trailing edge and into the pin fin bank. The structural support system reduces stresses proximate to the pin fins in the pin fin bank closest to the mid-chord region of the airfoil.



Inventors:
Metrisin, Joseph T. (Jupiter, FL, US)
Rogers, Friedrich (West Palm Beach, FL, US)
Rawlings, Christopher (Hobe Sound, FL, US)
Landis, Kenneth (Tequesta, FL, US)
Um, Jae (Orlando, FL, US)
Application Number:
11/415597
Publication Date:
11/08/2007
Filing Date:
05/02/2006
Assignee:
Siemens Power Generation, Inc.
Primary Class:
International Classes:
F01D5/18
View Patent Images:
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Primary Examiner:
WIEHE, NATHANIEL EDWARD
Attorney, Agent or Firm:
Siemens Corporation (Iselin, NJ, US)
Claims:
We claim:

1. A turbine airfoil, comprising: a generally elongated hollow airfoil formed from an outer wall, and having a leading edge, a trailing edge, a pressure side wall, a suction side wall positioned generally opposite from the pressure side wall; and a cooling system in the hollow airfoil, comprising: at least one trailing edge cooling channel positioned immediately adjacent to the trailing edge of the hollow airfoil; at least one cooling fluid supply channel in the hollow airfoil positioned proximate to the trailing edge cooling channel and extending between an inner surface of the pressure side wall and an inner surface of the suction side wall; a plurality of pin fins forming a pin fin bank extending between an inner surface of the pressure side wall and an inner surface of the suction side wall and positioned in the trailing edge cooling channel; and at least one chordal rib protruding from an inner surface of the outer wall of the hollow airfoil and extending in a general chordwise direction from a mid-chord region of the hollow airfoil toward the trailing edge and into the pin fin bank.

2. The turbine airfoil of claim 1, wherein the at least one chordal rib comprises a plurality of chordal ribs protruding from an inner surface of the outer wall of the hollow airfoil and extending in a general chordwise direction from a mid-chord region of the hollow airfoil toward the trailing edge and into the pin fin bank.

3. The turbine airfoil of claim 2, wherein the plurality of chordal ribs are positioned generally parallel to each other.

4. The turbine airfoil of claim 2, wherein the plurality of chordal ribs extend from an inner surface of the suction side wall to an inner surface of the pressure side wall in the pin fin bank.

5. The turbine airfoil of claim 1, further comprising a support rib extending between an inner surface of the pressure side wall and an inner surface of the suction side wall in the mid-chord region of the hollow airfoil and extending in a general spanwise direction to form a wall defining the at least one cooling fluid supply channel.

6. The turbine airfoil of claim 5, wherein at least one chordal rib extends from the support rib to the trailing edge of the hollow airfoil.

7. The turbine airfoil of claim 1, wherein the at least one chordal rib protruding from the outer wall of the hollow airfoil protrudes from the suction side wall.

8. The turbine airfoil of claim 1, wherein at least one chordal rib extends from an inner surface of the suction side wall to an inner surface of the pressure side wall in the pin fin bank.

9. The turbine airfoil of claim 1, wherein the at least one chordal rib protruding from an inner surface of the outer wall of the hollow airfoil comprises a plurality of protrusions extending from the at least one chordal rib within the pin fin bank.

10. The turbine airfoil of claim 1, wherein the at least one chordal rib protrudes from the inner surface of the outer wall into the at least one cooling fluid supply channel a distance between about two times and four times a thickness of the outer wall.

11. The turbine airfoil of claim 1, wherein plurality of pin fins forming the pin fin bank form a plurality of rows of pin fins that extend generally in a spanwise direction.

12. The turbine airfoil of claim 11, wherein the pin fins forming the rows extending in a generally spanwise direction are offset in a spanwise direction relative to pin fins in adjacent rows of pin fins.

13. The turbine airfoil of claim 1, further comprising a plurality of crossover ribs extending between the inner surface of the pressure side wall to the inner surface of the suction side wall and positioned in the cooling fluid supply channel in close proximity to the pin fin bank.

14. A turbine airfoil, comprising: a generally elongated hollow airfoil formed from an outer wall, and having a leading edge, a trailing edge, a pressure side wall, a suction side wall positioned generally opposite from the pressure side wall; and a cooling system in the hollow airfoil, comprising: at least one trailing edge cooling channel positioned immediately adjacent to the trailing edge of the hollow airfoil; at least one cooling fluid supply channel in the hollow airfoil positioned proximate to the trailing edge cooling channel and extending between an inner surface of the pressure side wall and an inner surface of the suction side wall; a support rib extending between an inner surface of the pressure side wall and an inner surface of the suction side wall in a mid-chord region of the hollow airfoil and extending in a general spanwise direction to form a wall defining the at least one fluid supply channel; a plurality of pin fins forming a pin fin bank extending between an inner surface of the pressure side wall to and an inner surface of the suction side wall and positioned in the trailing edge cooling channel; a plurality of chordal ribs protruding from an inner surface of the suction side wall of the hollow airfoil and extending in a general chordwise direction from the support rib in the mid-chord region of the hollow airfoil to the trailing edge and into the pin fin bank; and a plurality of crossover ribs extending between the inner surface of the pressure side wall to the inner surface of the suction side wall and positioned in the cooling fluid supply channel in close proximity to the pin fin bank.

15. The turbine airfoil of claim 14, wherein the plurality of chordal ribs extend from an inner surface of the suction side wall to an inner surface of the pressure side wall in the pin fin bank.

16. The turbine airfoil of claim 14, wherein the chordal ribs protruding from an inner surface of the outer wall of the hollow airfoil comprises a plurality of protrusions extending from ID sides of the chordal rib and from OD sides of the chordal ribs within the pin fin bank.

17. The turbine airfoil of claim 16, wherein the protrusions are aligned with rows of pin fins that extend in a spanwise direction forming the pin fin bank.

18. The turbine airfoil of claim 14, wherein plurality of pin fins forming the pin fin bank form a plurality of rows of pin fins that extend generally in a spanwise direction and wherein the pin fins forming the rows extending in a generally spanwise direction are offset in a spanwise direction relative to pin fins in adjacent rows of pin fins.

19. The turbine airfoil of claim 14, wherein the at least one chordal rib protrudes from the inner surface of the suction side wall into the at least one cooling fluid supply channel a distance between about two times and four times a thickness of the suction side wall.

20. A turbine airfoil, comprising: a generally elongated hollow airfoil formed from an outer wall, and having a leading edge, a trailing edge, a pressure side wall, a suction side wall positioned generally opposite from the pressure side wall; and a cooling system in the hollow airfoil, comprising: at least one trailing edge cooling channel positioned immediately adjacent to the trailing edge of the hollow airfoil; at least one cooling fluid supply channel in the hollow airfoil positioned proximate to the trailing edge cooling channel and extending between an inner surface of the pressure side wall and an inner surface of the suction side wall; a support rib extending between an inner surface of the pressure side wall and an inner surface of the suction side wall in a mid-chord region of the hollow airfoil and extending in a general spanwise direction to form a wall defining the at least one fluid supply channel; a plurality of pin fins forming a pin fin bank extending between an inner surface of the pressure side wall to and an inner surface of the suction side wall and positioned in the trailing edge cooling channel; and a plurality of crossover ribs extending between the inner surface of the pressure side wall to the inner surface of the suction side wall and positioned in the cooling fluid supply channel in close proximity to the pin fin bank.

Description:

FIELD OF THE INVENTION

This invention is directed generally to turbine airfoils, and more particularly to hollow turbine airfoils having cooling channels for passing fluids, such as air, to cool the airfoils.

BACKGROUND

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 assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine vane and blade assemblies to these high temperatures. As a result, turbine vanes and blades must be made of materials capable of withstanding such high temperatures. In addition, turbine vanes and blades often contain cooling systems for prolonging the life of the vanes and blades and reducing the likelihood of failure as a result of excessive temperatures.

Typically, turbine vanes are formed from an elongated portion forming a vane having one end configured to be coupled to a vane carrier and an opposite end configured to be movably coupled to an inner endwall. The vane is ordinarily composed of a leading edge, a trailing edge, a suction side, and a pressure side. The inner aspects of most turbine vanes typically contain an intricate maze of cooling circuits forming a cooling system. The cooling circuits in the vanes receive air from the compressor of the turbine engine and pass the air through the ends of the vane adapted to be coupled to the vane carrier. The cooling circuits often include multiple flow circuits that control metal temperature to ensure component durability and functionality. At least some of the air passing through these cooling circuits is exhausted through orifices in the leading edge, trailing edge, suction side, and pressure side of the vane.

Pin fin banks are commonly used within internal cooling chambers in turbine airfoils to increase heat transfer from the airfoil to the cooling fluids passing through internal cooling channels in the airfoil. In applications in which pin fin banks are positioned proximate to a trailing edge, high stresses typically occur immediately adjacent to the upstream row of pins closest to the leading edge as a result of thermal gradients between the suction side and pressure side walls. The high stresses caused by the thermal gradient often produce cracking in this region. Because of the thermal efficiencies created by using pin fin banks within turbine airfoils, a need exists for a turbine airfoil with a pin fin bank that is not as susceptible to damage from the thermal gradients between the suction side and pressure side of the turbine airfoil.

SUMMARY OF THE INVENTION

This invention relates to a turbine vane having an internal cooling system for removing heat from the turbine airfoil. The turbine airfoil may be formed from a generally elongated hollow airfoil having a leading edge, a trailing edge, a pressure side wall, a suction side wall and a cooling system. The cooling system may include chordal ribs for increasing the structural strength of a turbine airfoil having a plurality of pin fins forming a pin fin bank in a trailing edge cooling channel. The chordal ribs may extend in a general chordwise direction from a mid-chord region of the airfoil into the pin fin bank, thereby enhancing the structural stability of the turbine airfoil and reducing the likelihood of material failure at an intersection between the pin fins closest to the mid-chord region and suction side wall and in areas immediately surrounding the intersection.

The cooling system in the hollow airfoil may include at least one trailing edge cooling channel positioned immediately adjacent to the trailing edge of the hollow airfoil. The cooling system may also include at least one fluid supply channel in the hollow airfoil positioned proximate to the trailing edge cooling channel and extending between an inner surface of a pressure side wall and an inner surface of a suction side wall. A plurality of pin fins forming a pin fin bank may extend between the inner surface of the pressure side wall and the inner surface of the suction side wall. The pin fin bank may be positioned in the trailing edge cooling channel.

The cooling system may also include at least one chordal rib protruding from an inner surface of the outer wall of the hollow airfoil and extending in a general chordwise direction from a mid-chord region of the hollow airfoil to the trailing edge and into the pin fin bank. In at least one embodiment, the chordal ribs may extend from a support rib positioned in the mid-chord region, extend into the pin fin bank, and contact the trailing edge of the turbine airfoil. A portion of the chordal ribs in the pin fin bank may extend from an inner surface of the pressure side wall to the inner surface of the suction side wall. Such a configuration divides the trailing edge cooling channel into a plurality of cooling channels. The chordal ribs may include one or more protrusions extending from an ID surface or an OD surface, of both, on portions of the chordal ribs positioned within the pin fin bank. The protrusions may be aligned with rows of pin fins that extend generally in the spanwise direction. The protrusions may be positioned proximate to alternating rows of pin fins along the chordal ribs in the chordwise direction. The pin fins in the pin fin bank may be offset in the spanwise direction relative to pin fins in rows immediately adjacent to the pin fins.

The cooling system may also include one or more crossover ribs for increasing the structural integrity of the airfoil. In at least one embodiment, the cooling system may include a plurality of crossover ribs extending between the inner surface of the pressure side wall to the inner surface of the suction side wall and positioned in the cooling fluid supply channel in close proximity to the pin fin bank. The crossover ribs may have a cross-sectional area that is larger than the a cross-sectional area of a pin fin.

An advantage of this invention is that the chordal ribs provide a smooth load path between the pressure side wall and the suction side wall by reducing stress concentrations at the intersections between the pin fins and the pressure and suction side walls.

Another advantage of this invention is that the chordal ribs stiffen the entire section of the turbine airfoil proximate to the pin fin bank, thereby increasing the structural integrity of the airfoil. The chordal rib may create a contiguous and continuous panel stiffening rib that is tied to the pin bank.

Yet another advantage of this invention is that the chordal ribs functionally couple the pressure and suction sides together, thereby increasing the structural integrity of the airfoil.

Another advantage of this invention is that the chordal ribs prevent pressure induced bulging of the suction side wall.

These and other embodiments are described in more detail below.

BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings, which are incorporated in and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention.

FIG. 1 is a perspective view of a turbine airfoil having features according to the instant invention.

FIG. 2 is a cross-sectional view of the turbine airfoil shown in FIG. 1 taken along line 2-2.

FIG. 3 is a detailed view of a pin fin bank taken at detail line 3-3 in FIG. 2.

FIG. 4 is a partial, cross-sectional, perspective view of a cooling system in the turbine airfoil taken along line 4-4 in FIG. 1.

FIG. 5 is a partial, cross-sectional, perspective view of the cooling system shown in FIG. 4 as shown from a different perspective angle.

DETAILED DESCRIPTION OF THE INVENTION

As shown in FIGS. 1-5, this invention is directed to a turbine airfoil 10 for use in turbine engines that include a cooling system 12 in inner aspects of the turbine airfoil 10. The cooling system 12 may be used in any turbine vane or turbine blade. The cooling system 12 may include chordal ribs 34 for increasing the structural integrity of a turbine airfoil 12 having a plurality of pin fins 22 forming a pin fin bank 24 in a trailing edge cooling channel 14. The chordal ribs 34 may extend in a general chordwise direction from a mid-chord region 36 of the airfoil 10 into the pin fin bank 24, thereby enhancing the structural stability of the turbine airfoil 10 and reducing the likelihood of material failure at an intersection 54 between the pin fins 22 closest to the mid-chord region 36 and suction side wall 32. In at least one embodiment, the chordal ribs 34 extend in a general chordwise direction from a support rib 48 of the airfoil 10 into the pin fin bank 24.

As shown in FIG. 1, the turbine airfoil 10 may be formed from a generally elongated, hollow airfoil 18 having an outer surface 38 adapted for use, for example, in an axial flow turbine engine. The outer surface 38 may have a generally concave shaped portion forming a pressure side wall 28 and a generally convex shaped portion forming a suction side wall 32. The cooling system 12 may be used in a stationary turbine vane. In turbine blades, the turbine airfoil 10 may also include a second end 44 at one end. An outer endwall 42 may be coupled to an end opposite the tip section, and a second end 44 may be coupled to the outer endwall 42 and extend generally orthogonal to a longitudinal axis 46 of the airfoil 18. The airfoil may extend generally chordwise from a leading edge 50 to the trailing edge 16. In an embodiment, in which a turbine vane is used, as shown in FIGS. 1 and 2, the generally elongated airfoil 18 may include an outer endwall 42 adapted to be coupled to a hook attachment and may include a second end 44, opposite to the first end, that is adapted to be coupled to an inner endwall.

As shown in FIGS. 2 and 3, the cooling system 12 may include a trailing edge cooling channel 14 positioned immediately adjacent to a trailing edge 16 of the generally elongated, hollow airfoil 18. The trailing edge cooling channel 14 may extend between the second end 44 and the outer endwall 42 or may extend over any length of the distance between the second end 44 and the outer endwall 42. The trailing edge cooling channel 14 may extend from an inner surface 26 of the pressure side wall 28 to an inner surface 30 of the suction side wall 32 and may extend from the trailing edge 16 toward the mid-chord region 36. The configuration of the trailing edge cooling channel 14 may be any appropriate configuration for cooling internal aspects of the turbine airfoil 10.

The cooling system 12 may include a cooling fluid supply channel 20 in the hollow airfoil 18 positioned proximate to the trailing edge cooling channel 14. The cooling fluid supply channel 20 may extend between the inner surface 26 of the pressure side wall 28 and the inner surface 30 of the suction side wall 32. The cooling fluid supply channel 20 may extend between the second end 44 and the outer endwall 42 or may extend over any length of the distance between the second end 44 and the outer endwall 42. The cooling fluid supply channel 20 may be positioned in the mid-chord region 36 and may extend between the trailing edge cooling channel 14 and a support rib 48. The cooling fluid supply channel 20 may receive cooling fluids from any appropriate source, such as from a compressor (not shown) or from other cooling channels within the cooling system 12. The configuration of the cooling fluid supply channel 20 may be any appropriate configuration for cooling internal aspects of the turbine airfoil 10.

As shown in FIGS. 4 and 5, the cooling system 12 may also include a plurality of pin fins 22 forming a pin fin bank 24 extending between an inner surface 26 of a pressure side wall 28 to an inner surface 30 of a suction side wall 32. The pin fin bank 24 may be positioned in all regions of the trailing edge cooling channel 14 or only in portions of the trailing edge cooling channel 14. The pin fins 22 may have any appropriate configuration and cross-sectional shape. FIGS. 2 and 3 depict the pins as having a substantially circular cross-section; however, the cross-sectional shape of the pin fins 22 is not so limited. Rather, the pin fins 22 may have other appropriate cross-sectional shapes as well.

The cooling system 12 may include one or more chordal ribs 34 protruding from the inner surface 30 of the outer wall 32 of the hollow airfoil 18 and extending in a general chordwise direction from a mid-chord region 36 of the hollow airfoil 18 toward the trailing edge 16. The chordal ribs 34 may extend into the pin fin bank 24. In at least one embodiment, a plurality of chordal ribs 34 protrude from the inner surface 30 of the suction side wall 32. The chordal ribs 34 may protrude from the inner surface 30 of the suction side wall 32 into the cooling fluid supply channel 20 a distance between about two times and four times a thickness of the suction side wall 32. For instance, the height of the rib may be between about 0.20 inches to about 0.4 inches for a 0.10 thick suction side wall 32. In at least one embodiment, the chordal ribs 34 may be positioned generally parallel to each other. The chordal ribs 34 may extend from the support rib 48 and into the pin fin bank 24. In at least one embodiment, the chordal ribs 34 may extend from the support rib 48, into the pin fin bank 24, and contact the trailing edge 16. By extending the chordal ribs 34 from the support rib 48 and into the pin fin bank 24, the suction side wall 32 is supported at the intersection 54 between the pin fins 22 of the pin fin bank 24 closest to the mid-chord region 36. The chordal ribs 34 reduce the likelihood that the material at the intersection 54 and the suction side wall 32 immediately surrounding the intersection 54, will fail.

The portion of the chordal ribs 34 within the pin fin bank 24 may extend from an inner surface 26 of the pressure side wall 28 to an inner surface 30 of the suction side wall 32. Thus, the chordal ribs 34 within the pin fin bank 24 may divide the trailing edge cooling channel 14 into a plurality of trailing edge cooling channels 14, as shown in FIGS. 2 and 3. Additional support is provided to the region of the blade 18 between the trailing edge 16 and the mid-chord region 36 by extending the chordal ribs 34 between the pressure side wall 28 and the suction side wall 32 to further reduce the likelihood of premature failure of the turbine airfoil 10. The chordal ribs 34 provide a smooth load path between the pressure side wall 28 and the suction side wall 32 by reducing the stress concentrations at the intersection 54 between the pin fins 22 and the suction side wall 32. The chordal ribs 34 may also stiffen the pressure and suction side walls 28, 32 throughout portions of the airfoil 18 proximate to the pin fin bank 24.

As shown in FIG. 3, the chordal ribs 34 may include one or more protrusions 56. The protrusions 56 may extend from an ID surface 58 of the chordal rib 34 or from an OD surface 60 of the chordal rib 34, or both. The protrusions 56 may be aligned with rows 62 of pin fins 22 extending generally in a spanwise direction. The pin fins 22 within the rows 62 of pin fins 22 may be offset in the spanwise direction relative to pin fins 22 in adjacent rows 62 of pins fins 22, as shown in FIG. 3. In at least one embodiment, the protrusions 56 may be positioned adjacent to alternating rows 62 of pin fins 22, wherein the rows 62 of pin fins 22 extend in the spanwise direction, and the protrusions 56 are positioned proximate to alternating rows 62 along the chordal rib 34 in the chordwise direction.

The cooling system 12 may also include one or more crossover ribs 64, as shown in FIG. 4, for supporting the pressure and suction side walls 28, 32 to increase the structural integrity of the airfoil 18. In at least one embodiment, the cooling system 12 may include a plurality of crossover ribs 64 positioned immediately adjacent to the pin fin bank 24 closer to the leading edge 50, which in one embodiment is upstream from the pin fin bank 24. The crossover ribs 64 may be positioned in a row as shown in FIG. 4. The crossover ribs 64 may be sized larger than the pins 22. In particular, a crossover rib 64 may have a cross-sectional area that is between about 1.9 times and about 3.25 times larger than a cross-sectional area of a pin 22. The crossover ribs 64 may have a generally circular cross-section, as shown in FIG. 4, or may have any other appropriate shape. The crossover ribs 64 may extend from the inner surface 26 of the pressure side wall 28 to an inner surface 30 of the suction side wall 32. The crossover ribs 64 coupled the pressure side wall 28 and the suction side wall 32 together, thereby reducing the chordwise thermal strain in the pin fin bank 24 resulting from the thermal gradient that develops during operation between the pressure and suction side walls 28, 32.

During operation of a turbine engine in which the turbine airfoil 10 is employed, cooling fluids are passed into the cooling fluid supply channel 20 and are passed into the trailing edge cooling channel 14. The cooling fluids remove heat from the pressure and suction side walls 28, 32 and the pin fins 22. The cooling fluids are then exhausted through the trailing edge 16. The chordal ribs 34 provide enhanced structural support to the pressure side wall 28, the suction side wall 32, the trailing edge 16 and other regions of the turbine airfoil 10 during operation. The chordal ribs 34 distribute the stresses that form during operation to reduce the likelihood of premature failure of the turbine airfoil 10.

The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.