US20080085193A1

US20080085193A1 - Turbine airfoil cooling system with enhanced tip corner cooling channel

Info

Publication number: US20080085193A1
Application number: US11/543,523
Authority: US
Inventors: George Liang
Original assignee: Siemens Power Generations Inc
Current assignee: Siemens Energy Inc
Priority date: 2006-10-05
Filing date: 2006-10-05
Publication date: 2008-04-10

Abstract

A cooling system for a turbine airfoil of a turbine engine having a serpentine cooling channel with a portion positioned proximate to an intersection between a trailing edge and the tip section of the airfoil. The serpentine cooling channel may extend generally spanwise between a root and the tip section of the airfoil and may include a portion extending generally chordwise. The chordwise portion of the serpentine cooling channel may extend between the spanwise portion of the serpentine cooling channel and a trailing edge of the airfoil. The chordwise portion may include an exhaust orifice in the trailing edge and at a trailing edge turn in the serpentine cooling channel. The trailing edge turn provides a more efficient cooling channel for the cooling system which reduces the local temperature of the airfoil at the tip flag.

Description

FIELD OF THE INVENTION

This invention is directed generally to turbine airfoils, and more particularly to cooling systems in hollow turbine 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 blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine blades are formed from a root portion having a platform at one end and an elongated portion forming a blade that extends outwardly from the platform coupled to the root portion. The blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge. The inner aspects of most turbine blades typically contain an intricate maze of cooling channels forming a cooling system. The cooling channels in a blade 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 all aspects of the turbine blade at a relatively uniform temperature. However, centrifugal forces and air flow at boundary layers often prevent some areas of the turbine blade from being adequately cooled, which results in the formation of localized hot spots. Localized hot spots, depending on their location, can reduce the useful life of a turbine blade and can damage a turbine blade to an extent necessitating replacement of the blade. Thus, a need exists for a cooling system capable of providing sufficient cooling to turbine airfoils.

SUMMARY OF THE INVENTION

This invention relates to a turbine airfoil cooling system for a turbine airfoil used in turbine engines. In particular, the turbine airfoil cooling system includes a plurality of internal cavities positioned between outer walls of the turbine airfoil. The cavity may be formed from a mid-chord serpentine cooling channel positioned between a leading edge cooling channel and a trailing edge cooling channel. The mid-chord serpentine cooling channel may be formed from a spanwise portion extending in a generally spanwise direction between a root and a tip section and from a chordwise section extending in a generally chordwise direction between the spanwise section and a trailing edge of the airfoil. The mid-chord serpentine cooling channel may facilitate the efficient removal of heat from the airfoil, especially at the intersection between the tip section and the trailing edge.
The turbine airfoil cooling system may be positioned in a generally elongated, hollow airfoil having a leading edge, a trailing edge, a tip section at a first end, a root coupled to the airfoil at an end generally opposite to the first end for supporting the airfoil and for coupling the airfoil to a disc, and a cooling system formed from at least one cavity in the elongated, hollow airfoil. The cooling system may be formed from a serpentine cooling channel positioned in a mid-chord region of the generally elongated airfoil. The serpentine cooling channel may be formed from a first leg, a second leg, a third leg, a fourth leg, and a fifth leg. The first, fourth and fifth legs may be generally aligned with each other and may extend in a generally spanwise direction. The second and third legs may be generally aligned with each other and may extend in a generally chordwise direction proximate to an intersection between the trailing edge and the tip section of the generally elongated, hollow airfoil. The trailing edge turn may be positioned between the second and third legs and may include a trailing edge exhaust orifice in the trailing edge. A rib may be positioned in the trailing edge turn between the second and third legs. The rib may extend generally chordwise and may contact the trailing edge.
A trailing edge cooling channel may be in contact with the trailing edge and may extend from the root to the second leg of the serpentine cooling channel. One or more trailing edge exhaust orifices may be in fluid communication with the trailing edge cooling channel. One or more impingement ribs may be positioned in the trailing edge cooling channel that extends in a spanwise direction, wherein the impingement rib may include a plurality of impingement orifices. The cooling system may also include one or more leading edge cooling channels positioned proximate to the leading edge. One or more leading edge supply channels may be in fluid communication with the at least one leading edge cooling channel. A plurality of impingement orifices may be positioned in a rib separating the at least one leading edge supply channel from the at least one leading edge cooling channel.
The cooling system may include one or more trip strips in at least one of the legs of the serpentine cooling channel, the leading edge supply channel, the leading edge cooling channel, and the trailing edge cooling channel. The trip strips may protrude from an outer wall forming a pressure side of the generally elongated airfoil. In one embodiment, the trip strip may be formed by a plurality of trip strips protruding from the outer walls forming the pressure side and a suction side of the generally elongated airfoil.
An advantage of this invention is that the second and third legs of the serpentine cooling channel utilizes combined cooling fluids for the airfoil mid-body, blade tip section, and the tip flag, thereby increasing the total amount of cooling mass flow through the tip corner region and yields higher internal cooling capacity for the blade tip flag region at the intersection between the tip section and the trailing edge.
Another advantage of this invention is that a large portion of cooling fluids flows under the tip section, which provides additional convective cooling enhancement for the blade trailing edge tip section.
Still another advantage of this invention is that the pressure side tip section film cooling holes are angled in a direction opposite to the serpentine flow, thereby minimizing internal hole plugging problems.
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, commonly referred to as a filleted view, of the turbine airfoil shown in FIG. 1 taken along line 2-2.

FIG. 3 is a schematic diagram of the flow pattern of the cooling channels shown in FIG. 2.

DETAILED DESCRIPTION OF THE INVENTION

As shown in FIGS. 1-3, this invention is directed to a turbine airfoil cooling system 10 for a turbine airfoil 12 used in turbine engines. In particular, the turbine airfoil cooling system 10 includes a plurality of internal cavities 14, as shown in FIG. 2, positioned between outer walls 16 of the turbine airfoil 12. The cavity 14 may be formed from a mid-chord serpentine cooling channel 18 positioned between a leading edge cooling channel 20 and a trailing edge cooling channel 22. The mid-chord serpentine cooling channel 18 may be formed from a spanwise portion 24 extending in a generally spanwise direction between a root 26 and a tip section 28 and from a chordwise section 30 extending in a generally chordwise direction between the spanwise section 24 and a trailing edge 32 of the airfoil 12. The mid-chord serpentine cooling channel 18 may facilitate the efficient removal of heat from the airfoil 12, especially at the intersection 34 between the tip section 28 and the trailing edge 32.
As shown in FIGS. 1 and 2, the turbine airfoil 12 may be formed from a generally elongated, hollow airfoil 34 coupled to the root 26 at a platform 36. The turbine airfoil 12 may be formed from conventional metals or other acceptable materials. The generally elongated airfoil 34 may extend from the root 26 to the tip section 28 and may include a leading edge 38 and trailing edge 32. Airfoil 34 may have an outer wall 16 adapted for use, for example, in a first stage of an axial flow turbine engine. Outer wall 16 may form a generally concave shaped portion forming pressure side 40 and may form a generally convex shaped portion forming suction side 42. The cavity 14, as shown in FIG. 2, may be positioned in inner aspects of the airfoil 34 for directing one or more gases, which may include air received from a compressor (not shown), through the airfoil 34 to reduce the temperature of the airfoil 34.
The cooling system 10, as shown in FIGS. 2 and 3, may include a mid-chord serpentine cooling channel 18 forming the internal cavity 14. The mid-chord serpentine cooling channel 18 may be formed from a first leg 44, a second leg 46, a third leg 48, a fourth leg 50, and a fifth leg 52. The legs 44, 46, 48, 50 and 52 may be formed with internal ribs 54, 55, 57. The legs 44, 46, 48, 50 and 52 may extend from the outer wall 16 forming the pressure side 40 to the outer wall 16 forming a suction side 42. In other embodiments, the serpentine cooling channel 18 may be formed from more or less number of legs. The serpentine cooling channel 18 may extend from close proximity of the tip section 28 of the airfoil 34 to the root 26. In other embodiments, the serpentine cooling chamber 18 may have a shorter length.
The serpentine cooling chamber 18 may be in fluid communication with one or more cooling fluid supply channels 56 in the root 26. In particular, the first leg 44 may be in fluid communication with the supply channels 56 such that cooling fluids flow radially from the root 26 toward the tip section 28. The serpentine cooling chamber 18 may be positioned such that the first leg 44 extends in a generally spanwise direction and is positioned proximate to the trailing edge cooling channel 22. In at least one embodiment, the first, fourth and fifth legs 44, 50, 52 may be aligned with each other and may be positioned in a generally spanwise direction. The first, fourth and fifth legs 44, 50, 52 may form a spanwise portion 24 of the serpentine cooling chamber 18. The fourth leg 50 may be positioned between the first leg 44 and the fifth leg 52. The fourth and fifth legs 50, 52 may be separated by internal rib 57. The first and fourth legs 44, 50 may be separated by internal rib 55. The first and second legs 44, 46 may be in fluid communication through turn 72. The third and fourth legs 48, 50 may be in fluid communication through turn 74 and the fourth and fifth legs 50, 52 may be in fluid communication through turn 76. The turn 74 at the third and fourth legs 48, 50 may include a gap 86 that is used in the manufacturing process. The gap 86 is sealed with a plate 88.
As shown in FIG. 2, the second and third legs 46, 48 may be positioned proximate to the tip section 28 and may extend in a generally chordwise direction forming the chordwise portion 30. The second and third legs 46, 48 may be positioned between the trailing edge cooling channel 22 and the tip section 28. The second and third legs 46, 48 may be in fluid communication with each other through a trailing edge turn 58. The trailing edge turn 58 may be in contact with the trailing edge 32 and may include a trailing edge exhaust orifice 60 opening the serpentine cooling chamber 18 at the trailing edge 32. The second and third legs 46, 48 may be separated by the internal rib 54. The internal rib 54 may terminate in the chordwise direction toward the leading edge 38 from the trailing edge 32 to form the trailing edge turn 58. The size of the trailing edge turn 58 may be sized depending on the application of use of the turbine airfoil 12.
A rib 62 may be positioned in the trailing edge turn 58. The rib 62 may extend generally in line with the internal rib 54 separating the first and fourth legs 46, 50 and the second and third legs 46, 48. The rib 62 may extend generally in a chordwise direction and may extend from the outer wall 16 forming the pressure side 40 to the outer wall 16 forming the suction side 42.
As shown in FIG. 2, the turbine airfoil cooling system 10 may include a trailing edge cooling channel 64 in contact with the trailing edge 32 and extending from the root 26 to the second leg 46 of the serpentine cooling channel 18. The trailing edge cooling channel 64 may include one or more impingement ribs 66 extending in a generally spanwise direction. The impingement ribs 66 may include one or more impingement orifices 68. The impingement orifices 68 in adjacent impingement ribs 66 may be aligned or offset from each other. The trailing edge cooling channel 64 may also include a plurality of bleed off ribs 70 at the trailing edge 32. The bleed off ribs 70 may be staggered apart from each other to enable cooling fluids to be exhausted from the cooling system 10.
The turbine airfoil cooling system 10 may include one or more leading edge cooling channels 20 extending generally in a spanwise direction along the leading edge 38. A leading edge supply channel 78 may extend proximate to the leading edge cooling channel 20 and may be in fluid communication with the fluid supply channels 56. The leading edge supply channel 20 may be in fluid communication with the leading edge cooling channels 20 through one or more orifices 80.
The cooling system 10 may include one or more trip strips 82 for increasing the cooling capacity of the system 10. In at least one embodiment, the trip strips 82 may be positioned on inner surfaces 84 of the mid-chord serpentine cooling channel 18, the leading edge cooling channel 20, or the trailing edge cooling channel 22, or any combination thereof. The trips strips 82 may be positioned on the pressure side 40 or the suction side 42 of the cooling channels or both. The trip strips 82 may be positioned orthogonal to the flow of the cooling fluids through the cooling channels or may be positioned at an angle, as shown in FIG. 2.
The tip section 28 may include film cooling orifices 90 in fluid communication with the serpentine cooling channel 18 and the leading edge supply channel 78. The film cooling orifices 90 may be angled in a downstream direction to prevent the orifices from becoming clogged. The film cooling orifices 90 may also extend through the outer wall 16 forming the pressure and suction sides 40, 42.
During use, cooling fluids may flow into the cooling system 10 from a cooling fluid supply source (not shown) through the cooling fluid supply channel 56. In particular, the cooling fluids may enter into the first leg 44 of the serpentine cooling channel 18 and flow spanwise toward the tip section 28. The fluids may flow through the turn 72, through the second leg 46 and into the trailing edge turn 58 where the outer walls 16 proximate to the tip section 28 and the trailing edge 32 are cooled. Some of the cooling fluids may be exhausted through the trailing edge exhaust orifice 60 and the remainder of the cooling fluids may pass into the third leg 48. The cooling fluids may flow through the turn 74, into the fourth leg 50, through turn 76, into the fifth leg 52 flowing radially outwardly and be exhausted through film cooling orifices.
Cooling fluids may also flow from a cooling fluid supply channel 56 into the trailing edge cooling channel 64. The cooling fluids may flow through one or more impingement orifices 68 in the impingement rib 66 and past the bleed off ribs 70 where the cooling fluids are exhausted from the turbine airfoil 12. A portion of the cooling fluids may also be exhausted through film cooling orifices.
Cooling fluids may also flow from a cooling fluid supply channel 56 into the leading edge supply channel 78. Cooling fluids may flow from the leading edge supply channel 78 into the leading edge cooling channel 20 through one or more leading edge impingement orifices 80. The cooling fluids may be exhausted from the leading edge cooling channel 20 through film cooling orifices forming a showerhead.
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.

Claims

1. A turbine airfoil, comprising:

a generally elongated, hollow airfoil having a leading edge, a trailing edge, a tip section at a first end, a root coupled to the airfoil at an end generally opposite to the first end for supporting the airfoil and for coupling the airfoil to a disc, and a cooling system formed from at least one cavity in the elongated, hollow airfoil;

the cooling system comprising a serpentine cooling channel positioned in a mid-chord region of the generally elongated airfoil and being formed from a first leg, a second leg, a third leg, a fourth leg, and a fifth leg;

wherein the first, fourth and fifth legs are generally aligned with each other and extend in a generally spanwise direction;

wherein the second and third legs are generally aligned with each other and extend in a generally chordwise direction proximate to an intersection between the trailing edge and the tip section of the generally elongated, hollow airfoil; and

a trailing edge turn between the second and third legs including a trailing edge exhaust orifice in the trailing edge.

2. The turbine airfoil of claim 1, further comprising a rib positioned in the trailing edge turn between the second and third legs.

3. The turbine airfoil of claim 2, wherein the rib extends generally chordwise and contacts the trailing edge.

4. The turbine airfoil of claim 1, further comprising a trailing edge cooling channel in contact with the trailing edge and extending from the root to the second leg of the serpentine cooling channel.

5. The turbine airfoil of claim 4, further comprising at least one trailing edge exhaust orifice in fluid communication with the trailing edge cooling channel.

6. The turbine airfoil of claim 5, further comprising at least one impingement rib in the trailing edge cooling channel that extends in a spanwise direction, wherein the impingement rib includes a plurality of impingement orifices.

7. The turbine airfoil of claim 1, further comprising at least one leading edge cooling channel positioned proximate to the leading edge.

8. The turbine airfoil of claim 7, further comprising at least one leading edge supply channel in fluid communication with the at least one leading edge cooling channel.

9. The turbine airfoil of claim 8, further comprising a plurality of impingement orifices in a rib separating the at least one leading edge supply channel from the at least one leading edge cooling channel.

10. The turbine airfoil of claim 1, further comprising at least one trip strip in at least one of the legs of the serpentine cooling channel.

11. The turbine airfoil of claim 10, wherein the at least one trip strip protrudes from an outer wall forming a pressure side of the generally elongated airfoil.

12. The turbine airfoil of claim 11, wherein the at least one trip strip comprises a plurality of trip strips protruding from the outer walls forming the pressure side and a suction side of the generally elongated airfoil.

13. A turbine airfoil, comprising:

wherein the second and third legs are generally aligned with each other and extend in a generally chordwise direction between a trailing edge cooling channel and an intersection between the trailing edge and the tip section of the generally elongated, hollow airfoil;

a trailing edge trailing edge turn between the second and third legs including a trailing edge exhaust orifice in the trailing edge;

the trailing edge cooling channel in contact with the trailing edge and extending from the root to the second leg of the serpentine cooling channel and the trailing edge trailing edge turn; and

at least one leading edge cooling channel positioned proximate to the leading edge.

14. The turbine airfoil of claim 13, further comprising at least one trailing edge exhaust orifice in fluid communication with the trailing edge cooling channel.

15. The turbine airfoil of claim 14, further comprising at least one impingement rib in the trailing edge cooling channel that extends in a spanwise direction, wherein the impingement rib includes a plurality of impingement orifices.

16. The turbine airfoil of claim 15, further comprising at least one leading edge supply channel in fluid communication with the at least one leading edge cooling channel.

17. The turbine airfoil of claim 16, further comprising a plurality of impingement orifices in a rib separating the at least one leading edge supply channel from the at least one leading edge cooling channel.

18. The turbine airfoil of claim 13, further comprising at least one trip strip in at least one of the legs of the serpentine cooling channel, wherein the at least one trip strip comprises a plurality of trip strips protruding from the outer walls forming the pressure side and a suction side of the generally elongated airfoil.

19. The turbine airfoil of claim 13, further comprising a rib positioned in the trailing edge turn between the second and third legs.

20. The turbine airfoil of claim 19, wherein the rib extends generally chordwise and contacts the trailing edge.