METHOD FOR MODIFYING AN AIRFOIL SHROUD

Abstract

A turbine engine, bucket and method for modifying an airfoil shroud of turbine bucket is disclosed. A reference location is located in a second end edge of the airfoil shroud proximate a seal rail of the airfoil shroud. A relief cut is formed in the airfoil shroud to remove the reference location. Additionally, another reference location may be located in a first end edge of the airfoil shroud proximate the seal rail. Another relief cut may be formed in the airfoil shroud to remove the other reference location.

Description

BACKGROUND OF THE INVENTION

The subject matter disclosed herein relates to turbine engines. More particularly, the subject matter relates to modifying of turbine engine parts.

In a gas turbine engine, a compressor provides pressurized air to one or more combustors wherein the air is mixed with fuel and burned to generate hot combustion gas. These gases flow downstream to one or more turbines that extract energy therefrom to produce a mechanical energy output as well as power to drive the compressor. Over time, turbine parts, such as parts of the turbine, may experience fatigue, due to extreme conditions within the turbine, including high temperatures and pressures caused by flow of hot gas. In particular, certain turbine parts, such as buckets located on a turbine rotor, may experience fatigue that requires servicing or replacement.

In cases where reference locations in fatigued areas utilize welding or other heat-based operation, the repair process may further fatigue the local area. Thus, repair of some reference locations occurring due to wear and tear is not feasible. Replacement of these parts can be a costly, especially if fatigue in selected areas occurs in several parts, such as buckets on a rotor wheel.

BRIEF DESCRIPTION OF THE INVENTION

According to one aspect of the invention, a method is provided for modifying an airfoil shroud located at a tip of a blade of an airfoil, the airfoil shroud having a first end edge, a second end edge, a leading edge and a trailing edge, the method including: locating a reference location in the second end edge of the airfoil shroud, the reference location being proximate a seal rail extending circumferentially from a radially outer surface of the airfoil shroud; and forming a relief cut in the airfoil shroud to remove the reference location to thereby modify the airfoil shroud.

According to another aspect of the invention, a bucket is provided including: an airfoil having an airfoil axis; a shroud disposed at a tip of the airfoil, the shroud having a first end edge, a second end edge, a leading edge and a trailing edge; a seal rail extending circumferentially from a radially outer surface of the shroud; and a recess formed in the second end edge proximate the seal rail and on the leading edge of the airfoil.

According to another aspect of the invention, a turbine engine is provided including: a rotor; a bucket to be placed on the rotor, the bucket including: an airfoil having an airfoil axis; a shroud disposed at a tip of the airfoil, the shroud having a first end edge, a second end edge, a leading edge and a trailing edge; a seal rail extending circumferentially from a radially outer surface of the shroud; and a recess formed in the second end edge proximate the seal rail and the leading edge of the airfoil.

These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:

FIG. 1 is a schematic diagram of an embodiment of a gas turbine system;

FIG. 2 is a side view of an embodiment of a airfoil having a shroud;

FIG. 3 is a top view of the airfoil of FIG. 2;

FIG. 4 is a top view of an embodiment of a airfoil shroud having a flaw;

FIG. 5 is a top view of the airfoil shroud shown in FIG. 4 with one or more reliefs cut to repair the flaw;

FIG. 6 is an exemplary airfoil shroud from a third stage of a turbine engine; and

FIG. 7 is a flow chart of an exemplary process for modifying an airfoil shroud.

The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a schematic diagram of an embodiment of a gas turbine system 100. The system 100 includes a compressor 102, a combustor 104, a turbine 106, a shaft 108 and a fuel nozzle 110. In an embodiment, the system 100 may include a plurality of compressors 102, combustors 104, turbines 106, shafts 108 and fuel nozzles 110. As depicted, the compressor 102 and turbine 106 are coupled by the shaft 108. The shaft 108 may be a single shaft or a plurality of shaft segments coupled together to form shaft 108.

In an aspect, the combustor 104 uses liquid and/or gas fuel, such as natural gas or a hydrogen rich synthetic gas, to run the turbine engine. For example, fuel nozzles 110 are in fluid communication with a fuel supply 112 and pressurized air from the compressor 102. The fuel nozzles 110 create an air-fuel mix, and discharge the air-fuel mix into the combustor 104, thereby causing a combustion that creates a hot pressurized exhaust gas. The combustor 104 directs the hot pressurized exhaust gas through a transition piece into a rotor and stator assembly, causing turbine 106 to rotate as the gas exits nozzles and is directed onto the turbine buckets or blades. The rotation of the buckets coupled to the rotor in turbine 106 causes the shaft 108 to rotate, thereby compressing the air as it flows into the compressor 102.

In embodiments, one or more relief cuts may be formed in a shroud of an airfoil in the turbine engine. In an embodiment, the shroud is positioned on an airfoil such as a turbine bucket or a nozzle. The one or more relief cuts may be formed to modify the shroud and remove at least one of two reference locations in the airfoil shroud. In an embodiment, the at least one reference location is a flaw, such as a crack or missing material, that has been identified on the shroud. The at least one reference location may be caused by fatigue from exposure to extreme heat and pressure during turbine engine operation. In an embodiment, the relief cuts are formed without welding the shroud, thus reducing incidence of additional fatigue that may be introduced to the shroud by a welding process. In one embodiment, the relief cuts provide a structurally sound repair to the airfoil shroud to enable reuse and reinstallation of the airfoil following forming of the relief cuts. Accordingly, the repair process provides savings in time and costs when servicing the airfoil.

As used herein, “downstream” and “upstream” are terms that indicate a direction relative to the flow of working fluid through the turbine. As such, the term “downstream” refers to a direction that generally corresponds to the direction of the flow of working fluid, and the term “upstream” generally refers to the direction that is opposite of the direction of flow of working fluid. In addition, the terms “leading edge” and “trailing edge” indicate a position of a part relative to the flow of working fluid. Specifically, a leading edge of an airfoil encounters hot gas flow before a trailing edge of the airfoil. The term “radial” refers to movement or position perpendicular to an axis or center line of a reference part or assembly. It may be useful to describe parts that are at differing radial positions with regard to an axis. In this case, if a first component resides closer to the axis than a second component, it may be stated herein that the first component is “radially inward” of the second component. If, on the other hand, the first component resides further from the axis than the second component, it can be stated herein that the first component is “radially outward” or “outboard” of the second component. The term “axial” refers to movement or position parallel to an axis. Finally, the term “circumferential” refers to movement or position around an axis. Although the following discussion primarily focuses on gas turbines, the concepts discussed are not limited to gas turbines and may apply to any suitable rotating machinery, including steam turbines. Accordingly, the discussion herein is directed to gas turbine embodiments, but may apply to steam turbines and other turbomachinery.

FIG. 2 is a side view of a airfoil 200 according to an embodiment. FIG. 3 is a top view of the airfoil 200 shown in FIG. 3. In embodiments, a plurality of airfoils 200 is coupled to a rotor wheel in a turbine engine assembly, such as the turbine engine system 100. The airfoil 200 includes a blade 202. In an embodiment, the blade 202 converts the energy of a hot gas flow 206 into tangential motion of the bucket, which in turn rotates the rotor to which the bucket is attached. At the top of the blade 202, a seal rail 204 is provided to prevent the passage of hot gas flow 206 through a gap between the bucket tip and the inner surface of the surrounding stationary components (not shown). As depicted, the seal rail 204 extends circumferentially from a surface of a radially outer side 214 of a shroud 208 located at the bucket tip. As depicted, the shroud 208 includes the radially outer side 214 and a radially inner side 216. In an assembly of buckets on a rotor, the seal rail 204 extends circumferentially around a bucket row on the rotor, beyond the airfoil 200 sufficiently to line up with seal rails provided at the tip of adjacent buckets, effectively blocking flow from bypassing the bucket row so that airflow must be directed to the working length of the blade 202. During operation, the bucket row and rotor rotate about rotor axis 212. In addition, an airfoil axis 210 extends longitudinally through the blade 202.

In embodiments, the shroud 208 is a flat plate supported towards its center by the blade 202, where the shroud 208 is subject to high temperatures and centrifugal loads during turbine operation. As a result, portions of the shroud 208 may experience fatigue over time, where embodiments of the modifying process described herein repair fatigue, such as reference locations in the airfoil shroud.

FIG. 4 is a top view of an embodiment of an airfoil shroud 400 disposed at a tip of an airfoil as described above. The airfoil shroud 400 has a leading edge 402, a trailing edge 404, a first end edge 406 and a second end edge 408 defining the shroud. A seal rail 412 extends from a radially outer side 416 of the shroud in a circumferential direction from the first end edge 406 to the second end edge 408. In a bucket row assembly for a rotor, the first end edge 406 is configured to be placed adjacent the second end edge 408 of an adjacent airfoil shroud to provide a substantially continuous circumferential seal rail assembly in the turbine stage. The circumferential seal rail assembly blocks hot gas flow (e.g., 206) from bypassing the bucket row so that flow is directed along a working length of the bucket airfoil.

The seal rail 412 has fillets 414 on each side extending from the radially outer surface 416 to provide support for the seal rail 412. During operation of the turbine engine, fatigue caused by high pressures and temperatures can cause formation of a reference location B (410) that includes a deformation such as a crack and a reference location A (420) in the airfoil shroud 400 that includes a deformation such as a crack. In an embodiment, reference location B (410) is located on the trailing edge proximate the fillet 414 of seal rail 412 near the first end edge 406. Reference location A (420) is located on the leading edge near the fillet 414 of seal rail 412 near the second end edge 408. In general, the deformation at reference location A (420) may appear in the shroud after the deformation appears at reference location B (410) according to the normal use of the shroud 400. In cases where reference location A (420) is proximate structural regions, such as fillets 414, a relief cut (502, FIG. 5) may be used to repair and remove the reference location A (420), as described below. Similarly, when reference location B (410) is proximate structural regions, such as fillets 414, a relief cut (500, FIG. 5) may be used to repair and remove reference location B (410). The relief cuts 500 and 502 may be formed without performing a weld process on the shroud 400. In contrast, processes using welding to repair reference locations A and B may adversely affect material structural regions of the airfoil shroud 400, such as fillets 414.

Accordingly, FIG. 5 is a top view of the airfoil shroud 400 following a modifying of the airfoil shroud using the methods described herein. The method for modifying the airfoil shroud 400 includes locating reference location B (410) in the first end edge 406 of the shroud 400 and/or reference location A (420) in the second end edge 408 of the shroud 400. The modifying also includes forming a relief cut 500 in the first end edge 406 proximate the fillet 414 and/or forming relief cut 502 in the second end edge 508 proximate the fillet 414. In other embodiments, the relief cuts 500 and 502 may have any suitable geometry, such as a V-shape, parabolic, or polyhedron shape. The relief cut 500 may include a geometry that is the same as or different than a geometry of relief cut 502. In an embodiment, at least one of the relief cuts 500 and 502 may form an arc-shaped recess. The relief cuts 500 and 502 may be formed using any suitable process, such as machining or drilling, to remove material including the reference locations 410 and 420 from the airfoil shroud 400. The process for forming relief cut 500 may be the same as or different than the process for forming relief cut 502. In an embodiment, the airfoil shroud 400 is made from any suitable material, such as a steel alloy, stainless steel or other alloy.

In embodiments, the modifying process services or repairs the airfoil shroud 400 without a welding process, thus ensuring structural integrity is maintained in the region repaired. The structural integrity provided by the relief cuts 500 and 502 enables the airfoil shroud 400 to be reinstalled in the bucket row of the rotor and to withstand loads and stress caused by extreme temperatures and pressures. By forming the one or more relief cuts 500 and 502 as arc-shaped relief cuts, the resulting geometry, including the fillet 414, the first end edge 406 and the second end edge 408, maintains structural integrity to improve part life for the shroud, thus reducing operating costs for the turbine engine. In contrast, repair techniques that use a welding process may further fatigue the region being repaired. In some cases where welding is used for repair, welding may actually degrade the structural integrity of affected regions, thus leading to replacement of the entire airfoil and leading to increased operational costs. The service process utilizing relief cut 500 and relief cut 502 may be used to repair a reference location located in any suitable location, such as first end edge 406, second end edge 408, leading edge 402 and trailing edge 404. In various embodiments, the relief cuts 500 and 502 may remove a portion of the fillet 414 without resulting in significant structural losses. In other embodiments, at least one of the relief cuts 500 and 502 may be formed along a shroud edge and outside of the fillet 414. In cases where at least one of the relief cuts 500 and 502 forms an arc-shaped recess, a radius of the arc may vary depending on application needs.

FIG. 6 shows an exemplary airfoil shroud from a third stage of a turbine engine. The airfoil shroud 600 has a leading edge 602, a trailing edge 604, a first end edge 606 and a second end edge 608 defining the shroud. A seal rail 612 extends from a radially outer side 616 of the shroud in a circumferential direction from the first end edge 606 to the second end edge 608. The seal rail 412 has fillets 414 on each side extending from the radially outer surface 416 to provide support for the seal rail 412. Due to operation of the turbine engine, reference location A and/or reference location B may be formed. Reference locations A and B may include a flaw or deformation such as a crack, for example. Reference location A is located on the leading edge 602 near the fillet 614 of seal rail 612 near the second end edge 608. Reference location B is located on the trailing edge 604 proximate the fillet 614 of seal rail 612 near the first end edge 606. Reference locations A and B may form independently. Thus, reference location A may form while reference location B does not, and vice-versa. Additionally, both reference locations A and B may form together.

FIG. 6 further shows illustrative modifications that may be made to the airfoil shroud 600 using the methods described herein. Location A may be removed from the airfoil shroud 600 by machining a recess along relief cut 615. Relief cut 615 may have any suitable geometry, such as a V-shape, parabolic, or polyhedron shape. The relief cut 615 may form an arc-shaped recess. Similarly, location B may be removed from the airfoil shroud 600 by machining a recess along relief cut 617. Relief cut 617 may have any suitable geometry, such as a V-shape, parabolic, or polyhedron shape. The relief cut 617 may form an arc-shaped recess.

FIG. 7 is a flow chart 700 of an exemplary process for modifying an airfoil shroud. In block 702, a reference location (e.g., Location A), such as a crack, is located in a second end edge of an airfoil shroud, where the reference location is proximate a seal rail on the shroud. In embodiments, the reference location is on or proximate a fillet of the seal rail. In block 604, a relief cut is formed in the airfoil shroud surrounding the reference location, thus removing the reference location from the airfoil shroud. In block 606, another reference location (e.g. Location B), such as a crack, is located in a first end edge of an airfoil shroud, where the other reference location is proximate a seal rail on the shroud. In embodiments, the other reference location is on or proximate a fillet of the seal rail. In block 608, another relief cut is formed in the airfoil shroud surrounding the other reference location, thus removing the other reference location from the airfoil shroud. In block 610, the airfoil with the repaired shroud is replaced in a turbine engine. In an embodiment, the airfoil is placed in a second or third stage of the turbine engine. In embodiments, the modifying process may be complete after forming the relief cut or after forming the relief cut and the other relief cut, where the modifying does not include any welding of the shroud.

While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.

Claims

1. A method for modifying an airfoil shroud located at a tip of a blade of an airfoil, the airfoil shroud having a first end edge, a second end edge, a leading edge and a trailing edge, the method comprising: locating a reference location in the second end edge of the airfoil shroud, the reference location being proximate a seal rail extending circumferentially from a radially outer surface of the airfoil shroud; andforming a relief cut in the airfoil shroud to remove the reference location to thereby modify the airfoil shroud.

2. The method of claim 1, further comprising: locating another reference location in the first end edge of the airfoil shroud, the other reference location being proximate the seal rail; andforming another relief cut in the airfoil shroud to remove the other reference location.

3. The method of claim 2, further comprising locating the reference location proximate a fillet of the seal rail on the trailing edge of the airfoil shroud and locating the other reference location proximate the fillet on the leading edge of the airfoil shroud.

4. The method of claim 2, wherein forming the relief cut comprises forming a recess having a selected geometry in the second end edge of the airfoil shroud and forming the other relief cut comprises forming another recess of a selected geometry in the first end edge of the airfoil shroud.

5. The method of claim 4, wherein forming the recess and the other recess comprises forming at least one arc-shaped recess.

6. The method of claim 2, wherein forming the relief cut and the other relief cut further comprises machining at least one of the relief cut and the other relief cut.

7. The method of claim 1, comprising replacing the airfoil in at least one of: (i) a second stage of a turbine engine; and (ii) a third stage of a turbine engine after forming the relief cut.

8. A bucket comprising: an airfoil having an airfoil axis;a shroud disposed at a tip of the airfoil, the shroud having a first end edge, a second end edge, a leading edge and a trailing edge;a seal rail extending circumferentially from a radially outer surface of the shroud; anda recess formed in the second end edge proximate the seal rail and on the leading edge of the airfoil.

9. The bucket of claim 8, further comprising another recess formed in the first end edge proximate the seal rail and on the trailing edge of the airfoil.

10. The bucket of claim 10, wherein the recess comprises a relief cut of a selected geometry and wherein the other recess comprises a relief cut of a selected geometry.

11. The bucket of claim 10, wherein at least one of the geometry of the recess and the geometry of the other recess comprises an arc-shaped recess.

12. The bucket of claim 11, wherein the arc-shaped recess is formed by machining.

13. The bucket of claim 10, wherein the relief cut is formed to remove a reference location located in the second end edge proximate a fillet of the seal rail on the trailing edge and the other relief cut is formed to remove another reference location located in the first end edge proximate the fillet on the leading edge.

14. The bucket of claim 9, wherein the bucket is configured to be placed in one of: (i) a second stage of a turbine engine; and (ii) a third stage of a turbine engine.

15. A turbine engine comprising: a rotor;a bucket to be placed on the rotor, the bucket comprising:an airfoil having an airfoil axis;a shroud disposed at a tip of the airfoil, the shroud having a first end edge, a second end edge, a leading edge and a trailing edge;a seal rail extending circumferentially from a radially outer surface of the shroud; anda recess formed in the second end edge proximate the seal rail and the leading edge of the airfoil.

16. The turbine engine of claim 15, further comprising another recess formed in the first end edge proximate the seal rail and the trailing edge of the airfoil.

17. The turbine engine of claim 16, wherein at least one of the recess and the other recess comprises a relief cut.

18. The turbine engine of claim 17, wherein the relief cut is formed by at least one of machining or blending.

19. The turbine engine of claim 18, wherein the relief cut includes a relief cut formed to remove a reference location located in the second end edge proximate a fillet of the seal rail on the trailing edge and the other relief cut formed to remove another reference location located in the first end edge proximate the fillet on the leading edge.

20. The turbine engine of claim 15, wherein the bucket is located in at least one of: (i) a second stage of the turbine engine; and (ii) a third stage of the turbine engine.