WO2016133513A1

WO2016133513A1 - Turbine airfoil with a segmented internal wall

Info

Publication number: WO2016133513A1
Application number: PCT/US2015/016482
Authority: WO
Inventors: Evan C. LANDRUM; Jan H. Marsh; Paul A. SANDERS
Original assignee: Siemens Energy Inc
Current assignee: Siemens Energy Inc
Priority date: 2015-02-19
Filing date: 2015-02-19
Publication date: 2016-08-25
Anticipated expiration: 2017-08-19

Abstract

An airfoil (12) for a gas turbine engine in which the airfoil (12) includes an internal cooling system (10) formed in part by a segmented internal wall (54) that distributes mechanical loads is disclosed. The segmented internal wall (54) may include a plurality of connected segments (62). In one embodiment, at least a portion of the plurality of connected segments (62) may be non-parallel to a camber line (66) of the generally elongated hollow airfoil (32). In another embodiment, all of the plurality of connected segments (62) may be non-parallel to the camber line (66) of the generally elongated hollow airfoil (32). In a further embodiment, the plurality of connected segments (62) may be angled in an alternating fashion with respect to the camber line (66) of the generally elongated hollow airfoil (32).

Description

TURBINE AIRFOIL WITH A SEGMENTED INTERNAL WALL

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR

DEVELOPMENT

Development of this invention was supported in part by the United States

Department of Energy, Advanced Turbine Development Program, Contract No. DE- FC26-05NT42644. Accordingly, the United States Government may have certain rights in this invention. 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 airfoils (such as turbine vanes and blades) to high temperatures. As a result, turbine airfoils must be made of materials capable of withstanding such high temperatures, or must include cooling features to enable the component to survive in an environment which exceeds the capability of the material. As an example, typical airfoils may include internal cooling systems for reducing the temperature of the airfoils. Such typical internal cooling systems, however, may be deficient.

SUMMARY OF THE INVENTION

An airfoil for a gas turbine engine in which the airfoil includes an internal cooling system formed in part by a segmented internal wall that distributes mechanical loads is disclosed. The segmented internal wall may include a plurality of connected segments. In one embodiment, at least a portion of the plurality of connected segments may be non-parallel to a camber line of the generally elongated hollow airfoil. In another embodiment, all of the plurality of connected segments may be non-parallel to the camber line of the generally elongated hollow airfoil. In a further embodiment, the plurality of connected segments may be angled in an alternating fashion with respect to the camber line of the generally elongated hollow airfoil.

In at least one embodiment, a turbine airfoil may include a generally elongated hollow airfoil formed from an outer wall, and may have a leading edge, a trailing edge, a pressure side, and a suction side. The turbine airfoil may further include a cooling system positioned within an interior portion of the generally elongated hollow airfoil. The cooling system may be formed by the outer wall, a segmented internal wall, and a plurality of ribs positioned in-between the outer wall and the segmented internal wall. Cooling fluids may flow through the cooling system. The turbine airfoil may be a turbine blade, or a turbine vane.

The segmented internal wall may include a plurality of connected segments. At least a portion of the plurality of connected segments may be non-parallel to a camber line of the generally elongated hollow airfoil. Furthermore, all of the plurality of connected segments may be non-parallel to the camber line of the generally elongated hollow airfoil. The plurality of connected segments may be angled in an alternating fashion with respect to the camber line of the generally elongated hollow airfoil. Additionally, each two successive connected segments of the plurality of connected segments may be connected at an angle between 10 degrees and 170 degrees.

The cooling system may include an aft flowing serpentine cooling channel and further include a forward flowing serpentine cooling channel. The aft flowing serpentine cooling channel may be a five-pass aft flowing serpentine cooling channel. The aft flowing serpentine cooling channel may be an aft flowing spiral serpentine cooling channel, and the forward flowing serpentine cooling channel may be a forward flowing spiral serpentine cooling channel.

The plurality of ribs may connect the segmented internal wall to the outer wall. Additionally, at least a portion of the plurality of ribs may connect the segmented internal wall to either the pressure side of the outer wall or the suction side of the outer wall in an alternating fashion. In one embodiment, an advantage of the internal cooling system formed by the segmented internal wall is that the segmented internal wall may more easily deform relative to the outer wall. As such, the segmented internal wall may mitigate thermal/mechanical stresses caused by the outer wall being heated to a higher temperature than inner portions of the airfoil.

A further advantage of the internal cooling system formed by the segmented internal wall is that the segmented internal wall may reduce the cross-sectional area of the cooling cavities, and may also increase the length of the cooling cavities, in one embodiment. As such, the cooling fluids may more efficiently cool the airfoil, even for low flow designs.

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.

Figure 1 is a perspective view of a turbine airfoil according to one

embodiment.

Figure 2 is a cross-sectional view of an example of the turbine airfoil shown in Figure 1 taken along section line 2-2.

Figure 3 is a cross-sectional view of another example of the turbine airfoil shown in Figure 1 taken along section line 2-2.

DETAILED DESCRIPTION OF THE INVENTION

As shown in Figures 1-3, an airfoil 12 for a gas turbine engine in which the airfoil 12 includes an internal cooling system 10 formed by an outer wall 38, a segmented internal wall 54, and a plurality of ribs 58 positioned in-between the outer wall 38 and the segmented internal wall 54, is disclosed. The segmented internal wall 54 may include a plurality of connected segments 62. In one embodiment, at least a portion of the plurality of connected segments 62 may be non-parallel to a camber line 66 of the generally elongated hollow airfoil 32. In another embodiment, all of the plurality of connected segments 62 may be non-parallel to the camber line 66 of the generally elongated hollow airfoil 32. In a further embodiment, the plurality of connected segments 62 may be angled in an alternating fashion with respect to the camber line 66 of the generally elongated hollow airfoil 32.

Typically, a turbine airfoil (such as a turbine blade or a turbine vane) may be cooled by an internal cooling system that allows cooling fluids to pass through the turbine airfoil, cooling the outer wall of the turbine airfoil. Such internal cooling systems may be typically formed by a continuous internal wall that is parallel to a camber line of the turbine airfoil. However, this typical internal wall may be deficient because it may be subject to high stresses which may cause it to fail. For example, a typical internal wall may have a lower temperature than the outer wall of the airfoil. This temperature difference may cause the outer wall to deform at a different rate than the typical internal wall, which may subject the internal wall to high stress, and possible failure. Contrary to this, airfoil 12 of Figures 1 -3 may include an internal cooling system 10 that provides various advantages. For example, in one embodiment, an advantage of the internal cooling system 10 formed by the segmented internal wall 54 is that the segmented internal wall 54 may more easily deform relative to the outer wall 38. As such, the segmented internal wall 54 may mitigate thermal/mechanical stresses caused by the outer wall 32 being heated to a higher temperature than inner portions of the airfoil 12. As another example, in one embodiment, a further advantage of the internal cooling system 10 formed by the segmented internal wall 54 is that the segmented internal wall 54 may reduce the cross-sectional area of the cooling cavities, and may also increase the length of the cooling cavities. As such, the cooling fluids may more efficiently cool the airfoil 12, even for low flow designs.

Figure 1 illustrates a turbine airfoil 12 according to one embodiment. The airfoil 12 may be any type of turbine airfoil. For example, the airfoil 12 may be a turbine blade, a turbine vane, or any other turbine airfoil. As illustrated, the airfoil 12 is a turbine blade for a gas turbine engine. The airfoil 12 may include a generally elongated hollow airfoil 32 formed from an outer wall 38 adapted for use, for example, in a first stage of an axial flow turbine engine. The generally elongated hollow airfoil 32 may have a leading edge 46, a trailing edge 48, a generally concave shaped portion forming pressure side 40, and a generally convex shaped portion forming suction side 42. As illustrated, the generally elongated hollow airfoil 32 may be coupled to a root 34 at a platform 36. The root 34 may couple the airfoil 12 (such as a turbine blade) to a disc (not shown) of the turbine engine. In other

embodiments, the airfoil 12 may be a turbine vane with a first end coupled to the inner diameter of the turbine section of the turbine engine and a second end coupled to the outer diameter of the turbine section of the turbine engine.

A cavity 14 (shown in Figures 2-3) may be positioned in an inner portion of the airfoil 12 for directing one or more gases, which may include air received from a compressor (not shown), through the airfoil 12 and out one or more exhaust orifices 44 in the airfoil 12 to reduce the temperature of the airfoil 12. As shown in Figure 1 , the exhaust orifices 44 may be positioned in the leading edge 46, the trailing edge 48, the tip 50 in close proximity to the leading and trailing edges 46, 48, or any combination thereof, and have various configurations. The leading edge 46 may include a plurality of orifices 44 that collectively form a showerhead for cooling the leading edge 46 of the airfoil 12. Additionally, although Figure 1 illustrates orifices 44 as positioned at particular locations, the orifices 44 may be positioned anywhere along pressure side 40 and/or suction side 42. The cavity 14 may be arranged in various configurations and is not limited to a particular flow path.

A cooling system 10 may be located in cavity 14, as shown in Figures 2-3. As such, the cooling system 10 may be positioned within an interior portion of the generally elongated hollow airfoil 32. The cooling system 10 may be formed by a combination of the outer wall 38, a segmented internal wall 54, and a plurality of ribs 58 positioned in-between the outer wall 38 and the segmented internal wall 54. The segmented internal wall 54 may include a plurality of connected segments 62. The segmented internal wall 54 may have any number of connected segments 62. For example, the segmented internal wall 54 may have two connected segments 62, three connected segments 62, four connected segments 62, six connected segments 62, eight connected segments 62, or any other number of connected segments 62. A connected segment 62 may have any shape. For example, a connected segment 62 may be a straight segment, a curved segment, an irregular segment, any other shaped segment, or any combination of the preceding. A connected segment 62 may have any length and/or any width. In one embodiment, a portion of the connected segments 62 may be non- parallel to a camber line 66 halfway between the upper and lower surfaces of the generally elongated hollow airfoil 32. For example, one connected segment 62 may be non-parallel to the camber line 66, two connected segments 62 may be non- parallel to the camber line 66, three connected segments 62 may be non-parallel to the camber line 66, four connected segments 62 may be non-parallel to the camber line 66, or any other number of connected segments 62 may be non-parallel to the camber line 66. In one embodiment, all of the connected segments 62 may be non- parallel to the camber line 66. In another embodiment, half of the connected segments 62 (or any other portion of the connected segments 62, such as more than half or less than half) may be non-parallel to the camber line 66. A connected segment 62 may be non-parallel to the camber line 66 as a result of having an angle 68 with respect to the camber line 66. Angle 68 may be any angle. For example, angle 68 may be 5 - 85 degrees. In a further embodiment, the angle 68 of the connected segments 62 with respect to the camber line 66 may cause two successive connected segments 62 to be connected at an angle 70. Angle 70 may be any angle. For example, angle 70 may be 10 - 170 degrees.

As illustrated in Figures 2-3, the connected segments 62 may be angled in an alternating fashion with respect to the camber line 66. For example, a first connected segment 62 may be angled with respect to the camber line 66 so as to cause the first connected segment 62 to extend towards the pressure side 40 of the outer wall 38, while the next connected segment 62 may be angled with respect to the camber line 66 so as to cause the next connected segment 62 to extend towards the suction side 42 of the outer wall 38. In particular embodiments, this may cause a zig-zag type of shape for the segmented internal wall 54. Furthermore, this alternating fashion may continue along the entire length of the segmented internal wall 54, or only a portion of the segmented internal wall 54. In particular

embodiments, the segmented internal wall 54 may include one or more connected segments 62 that are not angled in an alternating fashion with respect to the camber line 66. For example, a first connected segment 62 may be angled with respect to the camber line 66 so as to cause the first connected segment 62 to extend towards the pressure side 40 of the outer wall 38, a next connected segment 62 (or more than one next connected segment 62) may also be angled so as to cause the next connected segment(s) 62 to extend towards the pressure side 40 of the outer wall 38 (or may be angled parallel to the camber line 66), and then the subsequent connected segment may be angled with respect to the camber line 66 so as to cause the subsequent connected segment 62 to extend towards the suction side 42 of the outer wall 38.

As is discussed above, the cooling system 10 may be formed by a

combination of the outer wall 38, a segmented internal wall 54, and a plurality of ribs 58 positioned in-between the outer wall 38 and the segmented internal wall 54. A rib 58 may refer to a segment (or other element) that connects the segmented internal wall 54 to the outer wall 38. For example, as illustrated in Figure 2, the ribs 58 may connect the segmented internal wall 54 to the outer wall 38. The airfoil 12 may include any number of ribs 58. For example, the airfoil 12 may include four ribs 58, five ribs 58, six ribs 58, eight ribs 58, nine ribs 58, ten ribs 58, or any other number of ribs 58. A rib 58 may have any shape. For example, a rib 58 may be a straight segment, a curved segment, an irregular segment, any other shaped segment, or any combination of the preceding. A rib 58 may have any length and/or any width. Furthermore, a rib 58 may be made of the same material as the segmented internal wall 54.

The ribs 58 may connect the segmented internal wall 54 to the outer wall 38 in any manner. For example, a rib 58 may connect the segmented internal wall 54 to both the pressure side 40 of the outer wall 38 and the suction side 42 of the outer wall 38. As another example, a rib 58 may connect the segmented internal wall 54 to either the pressure side 40 of the outer wall 38 or the suction side 42 of the outer wall 38. In such an example, such a connection may be done in an alternating fashion by at least a portion of the ribs 58, as is illustrated in Figure 2. The portion of the ribs 58 that connect the segmented internal wall 54 to either the pressure side 40 of the outer wall 38 or the suction side 42 of the outer wall 38 in an alternating fashion may include any number of ribs 58, such as two ribs 58, three ribs 58, four ribs 58, five ribs 58, six ribs 58, or any other number of ribs 58. In one embodiment, the portion of the ribs 58 that connect the segmented internal wall 54 to either the pressure side 40 of the outer wall 38 or the suction side 42 of the outer wall 38 in an alternating fashion may include all of the ribs 58. In a further embodiment, the portion of the ribs 58 that connect the segmented internal wall 54 to either the pressure side 40 of the outer wall 38 or the suction side 42 of the outer wall 38 in an alternating fashion may include half of the ribs 58 (or any other portion of the ribs 58, such as more than half or less than half).

In order to connect the segmented internal wall 54 to the outer wall 38, the ribs 58 may be connected to the connected segments 62. The ribs 58 may be connected to any location on the connected segments 62. For example, the ribs 58 may be connected to the end points of the connected segments 62, as is illustrated in Figure 2. As another example, the ribs 58 may be connected to middle points (or any other points) of the connected segments 62. As a further example, the ribs 58 may be connected to inflection points of curved connected segments 62. The ribs 58 may be connected to the connected segments 62 at an angle 76. Angle 76 may be any angle. For example, angle 76 may be 30 - 270 degrees. Furthermore, the ribs 58 may be connected to the outer wall 38 at an angle 78. Angle 78 may be any angle. For example, angle 78 may be 30 - 1 50 degrees.

Cooling system 10 may include one or more cooling channels that allow cooling fluids to flow through the cooling system 10 in order to cool the airfoil 12. The cooling channels may allow the cooling fluids to flow in the forward direction (i.e., a direction from the trailing edge 48 towards the leading edge 46), in the aft direction (i.e., a direction from the leading edge 46 towards the trailing edge 48), or both the forward direction and the aft direction. Furthermore, the one or more cooling channels may allow the cooling fluids to flow through all or a portion of the cooling system 10 in order to cool the airfoil 12.

As shown in Figures 2-3, cooling system 10 may include an aft flowing serpentine cooling channel 72 and a forward flowing serpentine channel 74. The aft flowing serpentine cooling channel 72 may extend from a position proximate the root 34 to the tip 50 of an airfoil 12 that is a turbine blade (or, for a turbine vane, from a position proximate the inner diameter of the turbine section to a position proximate the outer diameter of the turbine section). The aft flowing serpentine cooling channel 72 may be formed from at least a two pass serpentine cooling channel, and, in particular embodiments, may be a five-pass serpentine cooling channel or greater. As illustrated, the aft flowing serpentine cooling channel 72 may be an aft flowing spiral serpentine cooling channel having aft flowing cavities 82, 84, 86, 88, and 90 in contact with at least one of the pressure side 40 and the suction side 42. Aft flowing cavity 82 may be in communication with a first cooling fluid supply inlet (not shown), and may be configured to pass the cooling fluids through the aft flowing serpentine cooling channel 72 (via aft flowing cavities 84, 86, 88, and 90) to be exhausted from the airfoil through the trailing edge 42, and in at least one embodiment, through a trailing edge exhaust orifice 80.

The forward flowing serpentine cooling channel 74 may extend from a position proximate the root 34 to the tip 50 of an airfoil 12 that is a turbine blade (or, for a turbine vane, from a position proximate the inner diameter of the turbine section to a position proximate the outer diameter of the turbine section). The forward flowing serpentine cooling channel 74 may be formed from at least a two pass serpentine cooling channel, and, in particular embodiments, may be a three-pass serpentine cooling channel or greater. As illustrated in Figure 2, forward flowing serpentine cooling channel 74 may be a forward flowing spiral serpentine cooling channel having forward flowing cavities 94, 96, 98, and 100 in contact with at least one of the pressure side 40 and the suction side 42. Forward flowing cavity 94 may be in communication with a second cooling fluid supply inlet (not shown), and may be configured to pass the cooling fluids through the forward flowing cooling channel 74 (via forward flowing cavities 96, 98, and 100) to be exhausted from the airfoil through the leading edge 46, and in at least one embodiment, through leading edge orifices 44. In particular embodiments, cooling fluids flowing through forward flowing cavity 96 may be fed into forward flowing cavity 98 through one or more tangential holes 102, and the cooling fluids flowing through forward flowing cavity 98 may be fed into forward flowing cavity 100 through one or more tangential holes 104.

During use, cooling fluids may be passed from a cooling fluid supply (not shown), such as but not limited to, a compressor, to the airfoil 12. A first portion of the cooling fluids enter the aft flowing cavity 82 of the aft flowing serpentine cooling channel 72 via a first cooling fluid supply inlet (not shown). The first portion of the cooling fluids pass through the aft flowing serpentine cooling channel 72 in a spiral, serpentine manner, absorbing heat from the surfaces of the pressure side 40 and suction side 42 of the outer wall 38. For example, the first portion of cooling fluids may flow up (e.g., out of the page in Figure 2) aft flowing cavity 82 and pass over to aft flowing cavity 84, flow down (e.g., into the page in Figure 2) aft flowing cavity 84 and pass over to aft flowing cavity 86, flow up aft flowing cavity 86 and pass over to aft flowing cavity 88, flow down aft flowing cavity 88 and pass over to aft flowing cavity 90, and flow up aft flowing cavity 90 before being exhausted from the airfoil 12 through the trailing edge 48.

A second portion of the cooling fluids also enter the forward flowing cavity 94 of the forward flowing serpentine cooling channel 74 via a second cooling fluid supply inlet (not shown). The second portion of the cooling fluids pass through the forward flowing serpentine cooling channel 74 in a spiral, serpentine manner, absorbing heat from the surfaces of the pressure side 40 and suction side 42 of the outer wall 38. For example, the second portion of cooling fluids may flow up forward flowing cavity 94 and pass over to forward flowing cavity 96, flow down forward flowing cavity 96 and pass over to forward flowing cavity 98 through tangential holes 102. Furthermore, the cooling fluids flowing through forward flowing cavity 98 may be fed into forward flowing cavity 100 through one or more tangential holes 104. The second portion of cooling fluids may impinge on a backside surface of the leading edge 46 and may be exhausted through the orifices 44 forming a showerhead. As another example (such as shown in Figure 3), the second portion of the cooling fluids may flow up forward flowing cavity 94, pass over to forward flowing cavity 96, and flow down forward flowing cavity 96. Cooling fluids may be fed from cavity 96 into forward flowing cavity 98 through one or more cross over holes 102. Cooling fluids may impinge on a backside surface of the leading edge 46 and may be exhausted through the orifices 44 forming a showerhead.

Additionally, the cooling system 10 may include one or more additional elements and/or modifications. For example, the cooling system 10 may include one or more impingement orifices, one or more trailing edge pin fins, one or more blade tip holes, one or more rib turbulators, one or more film or gill holes, one or more refresher feeds, one or more fillets at the connection points between ribs 58 and the outer wall 38, any other elements for cooling airfoils 12, or any combination of the preceding. As another example, although the cooling fluids are described above as flowing through aft flowing serpentine cooling channel 72 and/or forward flowing serpentine cooling channel 74 in a spiral, serpentine manner, the cooling fluids may pass over and/or under any shared wall (such as a connected segment 62 or a rib 58) in the channel 72 and/or channel 74 while flowing through channel 72 and/or channel 74.

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

CLAIMS I claim:

1 . A turbine airfoil (12), characterized in that:

a generally elongated hollow airfoil (32) formed from an outer wall (38), and having a leading edge (46), a trailing edge (48), a pressure side (40), and a suction side (42); and

a cooling system (10) positioned within an interior portion of the generally elongated hollow airfoil (32), the cooling system (10) being formed by the outer wall (38), a segmented internal wall (54), and a plurality of ribs (58) positioned in-between the outer wall (38) and the segmented internal wall (54), wherein cooling fluids flow through the cooling system (10); and

wherein the segmented internal wall (54) includes a plurality of connected segments (62), wherein at least a portion of the plurality of connected segments (62) are non-parallel to a camber line (66) of the generally elongated hollow airfoil (32).

2. The turbine airfoil (12) of Claim 1 , characterized in that all of the plurality of connected segments (62) are non-parallel to the camber line (66) of the generally elongated hollow airfoil (32).

3. The turbine airfoil (12) of Claim 1 , characterized in that the plurality of connected segments (62) are angled in an alternating fashion with respect to the camber line (66) of the generally elongated hollow airfoil (32).

4. The turbine airfoil (12) of Claim 1 , characterized in that each two successive connected segments (62) of the plurality of connected segments (62) are connected at an angle (70) between 10 degrees and 170 degrees.

5. The turbine airfoil (12) of Claim 1 , characterized in that the turbine airfoil (12) is a turbine blade.

6. The turbine airfoil (12) of Claim 1 , characterized in that the turbine airfoil (12) is a turbine vane.

7. The turbine airfoil (12) of Claim 1 , characterized in that the cooling system (10) comprises an aft flowing serpentine cooling channel (72) and further comprises a forward flowing serpentine cooling channel (74).

8. The turbine airfoil (12) of Claim 7, characterized in that the aft flowing serpentine cooling channel (72) is an aft flowing spiral serpentine cooling channel (72) and wherein the forward flowing serpentine cooling channel (74) is a forward flowing spiral serpentine cooling channel (74).

9. The turbine airfoil (12) of Claim 7, characterized in that the aft flowing serpentine cooling channel (72) is a five-pass aft flowing serpentine cooling channel (72).

10. The turbine airfoil (12) of Claim 1 , characterized in that the plurality of ribs (58) connect the segmented internal wall (54) to the outer wall (38).

1 1 . The turbine airfoil (12) of Claim 10, characterized in that at least a portion of the plurality of ribs (58) connect the segmented internal wall (54) to either the pressure side (40) of the outer wall (38) or the suction side (42) of the outer wall (38) in an alternating fashion.